HANDBOOK OF COMPOSITES SECOND EDITION Edited by
S.T.Peters Process Research, Mountain View, Calfornia, USA
CHAPMAN & HALL
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London Weinheim . New York Tokyo Melbourne Madras
Published by Chapman & Hall, an imprint of Thomson Science, 2-6 Boundary Row, London SE18HN, UK Thomson Science, 2-6 Boundary Row, London SE18HN, UK Thomson Science, 115 Fifth Avenue, New York, NY 10003, USA Thomson Science, Suite 750,400 Market Street, Philadelphia, PA 19106, USA Thomson Science, Pappelallee 3,69469 Weinheim, Germany
First edition 1982 Second edition 1998
0 1998 Chapman & Hall Thomson Science is a division of International Thomson Publishing Typeset in 10/12 pt Palatino by GreenGate Publishing Services, Tonbridge, England Printed in Great Britain by Cambridge University Press ISBN 0 412 54020 7 All rights reserved. No part of this publication may be reproduced, stored in a retrieval system or transmitted in any form or by any means, electronic,,mechanical, photocopying, recording or otherwise, without the prior written permission of the publishers. Applications for permission should be addressed to the rights manager at the London address of the publisher. The publisher makes no representation, express or implied, with regard to the accuracy of the information contained in this book and cannot accept any legal responsibility or liability for any errors or omissions that may be made. A catalogue record for this book is available from the British Library
CONTRIBUTORS
SURESH G. ADVANI Department of Mechanical Engineering, University of Delaware, Spencer Laboratory, Newark, DE 19716, USA MAURICE E AMATEAU Applied Research Laboratory, Pennsylvania State University, PO Box 30, State College, PA 16804, USA
EVER J. BARBER0 315 Engineering Science Building, West Virginia University, Morgantown, WV 26506-6106, USA A.I. BEIL' Institute of Polymer Mechanics, Latvian Academy of Sciences, 23 Aizkraukles Street, Riga LV-1006, Latvia JEROME S. BERG True Temper Sports, 5421 Avenida Encinas, Suite G, Carlsbad, CA 92008, USA
KENNETH R. BERG Riggs Corporation, 837 Agate Street, Medford, OR 97501, USA LARS A. BERGLUND LuleH University of Technology, SE-97187 LuleH, Sweden D. BROWN Boeing Commercial Airplane Group, Douglas Products Division, Mail Stop D001-0018, Long Beach, CA 90846, USA JOHN D. BUCKLEY 23 East Governor Drive, Newport News, VA 23602, USA JERRY L. CADDEN C & S Technologies, 42759 Mountain Shadow, Murrieta, CA 92562, USA ZHONG CAI(deceased) 4180 Berkeley Creek Drive, Duluth, GA 30136, USA
x
Handbook of composites
FRANK A. CASSIS FAC Associates, 1150 N. Mountain, Suite 1028, Upland, CA 91786, USA
MIRIA M. FINCKENOR EH12 Bldg 4711, Marshall Space Flight Center, AL 35812, USA
LINDA L. CLEMENTS C & C Technologies, PO Box 1089, Dayton, NV 89403, USA
LIHWA FONG BLK G 5, Nanyang Avenue, Singapore 63616
DOUGLAS L. DENTON Chrysler Corporation, CIMS 482-00-13, 800 Chrysler Drive, Auburn Hills, MI 48326-2757, USA EDDY A. DERBY Composite Optics, 9617 Distribution Ave, San Diego, CA 92121, USA GEORGE W. DU Principal Engineer, 16331 Bay Vista Drive Cleanvater, FL 34620, USA HARRY W. DURSCH Boeing Defense and Space Group, PO Box 3999, Mail Stop 73-09, Seattle, WA 98124-2846, USA DON 0. EVANS Cincinnati Milacron, 4701 Marburg Avenue, Cincinnati, Ohio 45209, USA
HUGH H. GIBBS Polycomp Consulting, Inc., 25 Crestfield Road, Wilmington, DE 19810, USA TIMOTHY GUTOWSKI Department of Mechanical Engineering, Massachusetts Institute of Technology, Bldg 35-234, Cambridge, MA 02139, USA RICHARD N. HADCOCK 6 Sue Circle, Huntington, NY 11743, USA ENAMUL HAQUE Azdel, Inc., Technology Center, 658 Washburn Switch Road, Shelby, NC 28151-2284, USA
L.J. HART-SMITH Boeing Commercial Airplane Group, Douglas Products Division, Mail Stop D800-0019, 4000 Lakewood Boulevard, Long Beach, CA 90846, USA
Confributors xi JENNIFER HETH Cytec Fiberite, 501 W. Third Street, Winona, MN 55987-2854, USA
VALERY I. KOSTIKOV Niigrafit Institute, 2 Electrodonaya Street, Moscow, 111524, Russia
THOMAS S. JONES Industrial Quality, Inc., 640 E. Diamond Ave., Suite C, Gaithersburg, MD 20877, USA
GARY C. KRUMWEIDE Composite Optics, 9617 Distribution Avenue, San Diego, CA 92121, USA
THOMAS JUSKA Naval Surface Warfare Center, Carderock Division, Structures and Composites Department, Bethesda, MD 20084-5000, USA
V.L. KULAKOV Institute of Polymer Mechanics, Latvian Academy of Sciences, 23 Aizkraukles Street, Riga LV-1006, Latvia
JOHN T. KANNE 2201 Johnson Road, Memphis, TN 38139, USA
KHALID LAFDI Center for Advanced Friction Studies, Southern Illinois University at Carbondale, Carbondale, IL 62901-4343, USA
HARRY S. KATZ Utility Development Corporation, 112 Naylon Avenue, Livingston, NJ 07039, USA
CHRISTY KIRCHNER LAPP 1412 Bellingham Way, Sunnyvale, CA 94087, USA
V.S. KILIN Niigrafit Institute, 2 Electrodonaya Street, Moscow, 111524, Russia FRANK K. KO Drexel University, Fibrous Materials Research Laboratory 27-439, Philadelphia, PA 19104, USA KENT E. KOHKONEN Brigham Young University, 435 CTB Technology Department, Provo, UT 84602, USA
ROBERT A. LATOUR Clemson University, Clemson, SC 29634, USA BURR L. LEACH Cambridge Industries, 1700 Factory Avenue, Marion, IN 46952, USA STEWART N. LOUD Composites Worldwide Inc., 991 Lomas Santa Fe Drive, C469, Solana Beach, CA 92075-2125, USA
xii Handbook of composites VICKI P. MCCONNELL Ray Publishing, Independence Street, Suite 270, Wheat Ridge, CO 80033, USA
NITIN POTDAR Brigham Young University, 435 CTB Technology Department, Provo, UT 84602, USA
ANDREW C. MARSHALL Marshall Consulting, 720 Appaloosa Drive, Walnut Creek, CA 94596, USA
KENNETH REIFSNIDER Virginia Polytechnic Institute and State University, Patton Hall 120, Blacksburg, VA 24061-0219, USA
ANTHONY MARZULLO 39 Harold Street, COSCob, CT 06807-2132, USA
THEODORE J. REINHART 345 Forrer Boulevard, Dayton, OH 45419-3238, USA
DONALD W. OPLINGER Federal Aviation Administration, Wm. J. Hughes Technical Center AAR-431, Atlantic City, International Airport, NJ 08405, USA
PAUL E SADESKY C & S Technologies, 23547 Mountain Court, Murrieta, CA 92562, USA
HARRY E. PEBLY 198 Center Grove Road, Randolph, NJ 07869, USA LYNN S. PENN Department of Chemical and Materials Engineering, 177 Anderson Hall, University of Kentucky, Lexington, KY 40506-0046, USA S.T. PETERS Process Research, 925 Sladky Avenue, Mountain View, CA 94040-3625, USA
FRANK J. SCHWAN 36671 Montecito Drive, Fremont, CA 94536, USA ANTON L. SEIDL 18941 Mellon Drive, Saratoga, CA 95070, USA JOCELYN M. SENG Owens Corning Science and Technology Center, 2790 Columbus Road, Granville, OH 43023-1200, USA SHALABY W. SHALABY Clemson University, 301 Rhodes Res., Clemson, SC 29634, USA
Contributors xiii DAVID A. SHIMP PO Box 974, Prospect, KY 40059, USA DONALD R. SIDWELL 44609 Grove Lane, Lancaster, CA 93534-2833, USA BRIAN E. SPENCER Spencer Composite Corporation, 3220 Superior Street, PO Box 4377, Lincoln, NE 68504-0377, USA ROBERT C. TALBOT 7199 Lorine Court, Columbus, OH 43235-5125, USA YU.M. TARNOPOL'SKII Institute of Polymer Mechanics, Latvian Academy of Sciences, 23 Aizkraukles Street, Riga LV-1006, Latvia R.C. TENNYSON University of Toronto, Institute for Aerospace Studies, 4925 Dufferin Street, Downsview, Ontario, Canada M3H 5T6 JAMES L. THRONE Shenvood Technologies, Inc., 158 Brookside Boulevard, Hinckley, OH 44233-9676, USA FRANK TRACESKI Department of Defense, 5203 Leesburg Pike Suite 1403, Falls Church, VA 22041, USA
WAYNE C. TUCKER Naval Undersea Warfare Center, PO Box 86, Exeter, RI 02822, USA V. V. VASILIEV Moscow State University, 14-1-110 Podolskih Kursantov Street, Moscow 113545, Russia DENNIS J. VAUGHAN 146 Longview Drive, Anderson, SC 29621, USA H. WANG Department of Chemical and Materials Engineering, 177 Anderson Hall, University of Kentucky, Lexington, KY 40506-0046, USA ANN E WHITAKER EHOl Bldg 4612, Marshall Space Flight Center, AL35812, USA BRIAN A. WILSON Wilson Composite Group, 6611 Folsom-Auburn Road, Suite C, Folsom, CA 95630, USA
S. WONG Boeing Commercial Airplane Group, Douglas Products Division, Mail Stop D001-0018, Long Beach, CA 90846, USA
xiv Handbook of composites MAURICE A. WRIGHT Center for Advanced Friction Studies, Southern Illinois University at Carbondale, Carbondale, IL 62901-4343, USA
PHILIP R. YOUNG Emory & Henry College, Department of Chemistry, Emory, VA 24327, USA
ABOUT THE EDITOR
S.T. Peters was previously a fellow engineer with Westinghouse Electric Corporation, Marine Division prior to devoting full time to composite and materials and processing consulting for his own company, Process Research, in Mountain View, CA. He has written many articles on composites and filament winding, a book on filament winding, edited one previous book and holds several patents on winding techniques and composite joints.
He is a private consultant with worldwide clients and has presented tutorials on composites to many audiences, including the US Navy and NASA, several technical societies and two universities. He is a licensed professional engineer in the state of California, a member of ASM, and the composites division of SME and has been elected a fellow of SAMPE.
ACKNOWLEDGEMENTS
As with any large undertaking there is a supporting group of people without whose help the objective would not be met. I wish to acknowledge my wife, Lynn, for her help in deciphering and rewriting some of the articles and for enduing my sometimes uncivil
approach to resolving problems. Thanks also go to Mr Frank Heil and Dr Alvin Nakagawa of Westinghouse Electric, Marine Division (now Norton Grumman) for their editorial and review help. I also wish to thank Dr Linda Clements for her advice and support.
PREFACE
Today, fiber reinforced composites are in use in a variety of structures, ranging from spacecraft and aircraft to buildings and bridges. This wide use of composites has been facilitated by the introduction of new materials, improvements in manufacturing processes and developments of new analytical and testing methods. Unfortunately, information on these topics is scattered in journal articles, in conference and symposium proceedings, in workshop notes, and in government and company reports. This proliferation of the source material, coupled with the fact that some of the relevant publications are hard to find or are restricted, makes it difficult to identify and obtain the up-to-date knowledge needed to utilize composites to their full advantage. This book intends to overcome these difficulties by presenting, in a single volume, many of the recent advances in the field of composite materials. The main focus of this book is on polymeric matrix, metal matrix, and ceramic matrix composites. The book treats a wide range of subjects. The topics, presented in 49 chapters and two appendices include: 0
overview of composite material systems and products;
0 0
0 0 0 0 0 0
0 0
properties of different component (fiber, matrix, filler) materials; manufacturing techniques; analysis and design; testing; mechanically fastened and bonded joints; repair; damage tolerance; environmental effects; health, safety, reuse, and disposal; applications in: aircraft and spacecraft; land transportation; marine environments; biotechnology; construction and infrastructure; sporting goods.
Each chapter, written by a recognized expert, is self-contained, and contains many of the 'state-of-the-art' techniques required for practical applications of composites. Thus, this book should serve as a useful source of information for practicing engineers and specialists, as well as for workers new to this field. George S. Springer
CONTENTS
Contributors
ix
Preface
xv
About the editor
xvi
Foreword
xvii
Acknowledgements
xviii
Introduction, composite basics and road map S.T. Peters 1 Overview of composite materials
1
21
Theodore J. Reinhart PART ONE: BASIC MATERIALS Polymeric matrix systems 2 Polyester and vinyl ester resin Frank A. Cassis and Robert C. Talbot
34
3 Epoxyresins L.S. Penn and H . Wang
48
4 High temperature resins
75
Hugh H . Gibbs 5 Speciality matrix resins David A . Shimp 6 Thermoplastic resins Lars A. Berglund
99 115
Reinforcements and composites 7 Fiberglass reinforcement Dennis J. Vaughan
131
vi Handbook of composites 8 Boron, high silica, quartz and ceramic fibers Anthony Marzullo
156
9 Carbon fibers Khalid Lafdi and Maurice A. Wright
169
10 Organic fibers Linda L. Clements
202
11 Particulate fillers Harry S. Katz
242
12 Sandwich construction Andrew C. Marshall
254
13 Metal matrix composites V l . Kostikov and V S . Kilin
29 1
14 Ceramic composites M.E Amateau
307
15 Carbon-carbon composites John D. Buckley
333
PART TWO: PROCESSING METHODS General composites and reinforced plastics 16 Hand lay-up and bag molding D.R. Sidwell
352
17 Matched metal compression molding of polymer composites Enamul Haque and Burr (Bud) L. Leach
378
18 Textile preforming Frank K. KO and George W. Du
397
19 Table rolling of composite tubes John T. Kanne and Jerome S. Berg
425
20 Resin transfer molding Lihwa Fong and S.G. Advani
433
21 Filament winding Yu.M. Tarnopol’skii, S.T. Peters, A.I. Beil’
456
22 Fiber placement Don 0. Evans
476
23 Pultrusion Brian A. Wilson
488
24 Processing thermoplastic composites James L. Throne
525
Contents vii Advanced composites 25 Tooling for composites Jerry L. Cadden and Paul F. Sadesky
556
26 Consolidation techniques and cure control Zhong Cui and Timothy Gutowski
576
27 Composite machining Kent E. Kohkonen and Nitin Potdar
596
28 Mechanical fastening and adhesive bonding D. W. Oplinger
610
29 Surface preparations for ensuring that the glue will stick in bonded composite structures L.J. Hart-Smith, D. Brown and S. Wong
667
PART THREE: DESIGN AND ANALYSIS 30 Laminate design Jocelyn M . Seng
686
31 Design of structure with composites F.J. Schwan
709
32 Analysis methods V.V. Vasiliev
736
33 Design allowables substantiation Christy Kirchner Lapp
758
34 Mechanical tests Yu.M. Tarnopol'skii and V.L. Kulakov
778
PART FOUR. ENVIRONMENTAL EFFECTS 35 Durability and damage tolerance of fibrous composite systems Ken Reifsnider
794
36 Environmental effects on composites A n n F. Whitaker, Miria M . Finckenor, Harry W. Dursch, R.C. Tennyson and Philip R. Young
810
37 Safety and health issues
822
Jennifer A. Heth 38 Nondestructive evaluation methods for composites Thomas S. Jones
838
39 Repair aspects of composite and adhesively bonded aircraft structures Anton L. Seidl
857
40 Reuse and disposal Harry E , Pebly
883
viii Handbook of composites
PART FIVE APPLICATIONS 41 Land transportation applications Douglas L. Denton
905
42 Marine applications Wayne C. Tucker and Thomas Juska
916
43 Commercial and industrial applications of composites Stewart N. Loud
931
44 Composite biomaterials Shalaby W. Shalaby and Robert A. Latour
957
45 Scientificapplications of composites Vicki I? McConnell
967
46 Construction Ever J. Barber0
982
47 Aerospace equipment and instrument structure G a y C. Krumweide and Eddy A. Derby
1004
48 Aircraft applications Richard N. Hadcock
1022
49 Composites in the sporting goods industry Brian E. Spencer
1044
APPENDICES Appendix A Typical properties for advanced composites Kenneth R. Berg
1053
Appendix B Specifications and standards for polymer composites Frank T. Traceski
1059
Index
1069
INTRODUCTION, COMPOSITE BASICS AND ROAD MAP* S.T. Peters
This is an introduction to composites and will encourage the reader to obtain more information. Only the basic concepts will be covered here; reference will be made to the chapters in the book that expand or follow up and elaborate on these basics. The reader will see that the subjects of this book cover the spectrum of composites and range from the basic and simple to the complex. Thus, there are complicated equations because they are the tools that are used every day to describe real structures; and there will also be the more general, less complicated approaches that are limited in analysis power. These chapters have been developed by the most knowledgeable composite professionals in the world; a blend of academicians and the engineers who fabricate real composite structures. Modern structural composites, frequently referred to as ’Advanced Composites’, are a blend of two or more components, one of which is made up of stiff, long fibers, and the other, a binder or ’matrix’ which holds the fibers in place. The fibers are strong and stiff relative to the matrix and are generally orthotropic (having different properties in two different directions). The fiber, for advanced structural composites, is long, with length to diameter ratios of over 100. The fiber’s strength and stiffness are usually much greater, perhaps several times more, than the matrix material. The matrix material can by polymeric (e.g. polyester resins, epoxies), Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
metallic, ceramic or carbon. When the fiber and the matrix are joined to form a composite they retain their individual identities and both directly influence the composite’s final properties. The resulting composite will generally be composed of layers (laminae) of the fibers and matrix stacked to achieve the desired properties in one or more directions. The high strength or stiffness to weight ratios of advanced composites are well known, but there are other advantages also (Table 1.1). These advantages translate not only into aircraft, but into everyday activities, such as longer drives with a graphite-shafted golf club (because more of the mass is concentrated at the clubhead) or less fatigue and pain because a graphite composite tennis racquet has mherent damping. Generally, the advantages accrue for any fiber/composite combination and disadvantages are more obvious with some. These advantages have now resulted in many more reasons for composite use as shown in Table 1.2. Proper design and material selection can circumvent many of the disadvantages. 1.1 MATERIAL SYSTEMS
An advanced composite laminate can be tailored so that the directional dependence of strength and stiffness matches that of the loading environment. To do that, layers of unidirectional material called laminae are ori* This chapter has been adapted from S.T. Peters, in Handbook of Plastics Elastomers and Composites, 3rd edn, (ed. C.A. Harper). McGraw-Hill, New York, 1996, and is used with permission of the McGraw-Hill companies.
2 Introduction, composite basics and road map Table 1.1 Advantages/disadvantages of advanced composites Advantages
Disadvantages
Weight reduction High strength or stiffness to weight ratio
Cost of raw materials and fabrication
Tailorable properties Can tailor strength or stiffness to be in the load direction
Transverse properties may be weak
Redundant load paths (fiber to fiber)
Matrix is weak, low toughness
Longer life (no corrosion)
Reuse and disposal may be difficult
Lower manufacturing costs because of less part count
Difficult to attach
Inherent damping
Analysis is difficult
Increased (or decreased) thermal or electrical conductivity
Matrix subject to environmentaldegradation
___-
Carbon/graphite fibers (Chapter 9) have demonstrated the widest variety of strengths and modulii and have the greatest number of suppliers. The fibers begin as an organic fiber, rayon, polyacrylonitrile or pitch which is called the precursor. The precursor is then stretched, oxidized, carbonized and graphitized. There are many ways to produce these fibers, but the relative amount of exposure at temperatures from 2500-3000°C results in greater or less graphitization of the fiber. Higher degrees of graphitization usually result in a stiffer fiber (higher modulus) with 0 fiberglass; greater electrical and thermal conductivities 0 graphite; and usually higher cost. 0 aramid; The organic fiber Kevlar 49, (Chapter 10) 0 polyethylene; also called aramid, essentially revolutionized 0 boron; pressure vessel technology because of its great 0 silicon carbide; tensile strength and consistency coupled with 0 silicon nitride, silica, alumina, alumina silica. low density, resulting in much more weight The advantages of fiberglass (Chapter 7) are its effective designs for rocket motors. Aramid high tensile strength and strain to failure, but composites are still widely used for pressure heat and fire resistance, chemical resistance, vessels but have been largely supplanted by moisture resistance and thermal and electrical the very high strength graphite fibers. Aramid properties are also cited as reasons for its use. composites have relatively poor shear and It is by far the most widely used fiber, primar- compression properties; careful design is ily because of its low cost; but its mechanical requires for their use in structural applications properties are not comparable with other that involve bending or compression. structural fibers.
ented to satisfy the loading requirements. These laminae contain fibers and a matrix. Because of the use of directional laminae, the tensile, flexural and torsional shear properties of a structure can be disassociated from one another to some extent and a golf shaft, for example, can be changed in torsional stiffness without changing the flexural or tensile stiffness. Fibers can be of the same material within a lamina or several fibers mixed (hybrid). The common commercially available fibers are as follows:
Material systems 3 Table 1.2 The reasons for using composites
Reason for use
Material selected
___
Appl ica t ion/driver
Lighter, stiffer stronger
Boron, all carbodgraphites, some aramid
Military aircraft, better performance Commercial aircraft, operating costs
Controlled or zero thermal expansion
Very high modulus carbon/graphite
Spacecraft with high positional accuracy requirements for optical sensors
Environmental resistance
Fiberglass, vinyl esters, bisphenol A fumarates, chlorendic resins
Tanks and piping, corrosion resistance to industrial chemicals, crude oil, gasoline at elevated temperatures
Lower inertia, faster startups, less deflection
High strength carbon/graphite, epoxy
Industrial rolls, for paper, films
Lightweight, damage tolerance
High strength carbon/graphite, CNG tanks for ’green’cars, trucks fiberglass, (hybrids), epoxy and busses to reduce environmental pollution
More reproducible complex surfaces
High strength or high modulus carbon graphite/ epoxy
High-speed aircraft. Metal skins cannot be formed accurately
Less pain and fatigue
Carbon/graphite/epoxy
Tennis, squash and racquetball racquets. Metallic racquets are no longer available
Reduces logging in ‘old growth’ forests
Aramid, carbon/graphite
Laminated ‘new’ growth wooden support beams with high modulus fibers incorporated
Reduces need for intermediate support and resists constant 100% humidity atmosphere
High strength carbon/graphite-epoxy
Cooling tower driveshafts
Tailorability of bending and twisting response
Carbon/graphite-epoxy
Golf shafts, fishing rods
Transparency to radiation
Carbon/ graphite-epoxy
X-ray tables
Crashworthiness
Carbon/ graphite-epoxy
Racing cars
Higher natural frequency, lighter Carbon/ graphite-epoxy
Automotive and industrial driveshafts
Water resistance
Fiberglass (woven fabric), polyester or isopolyester
Commercial boats
Carbon/graphite, fiberglass
Freeway support structure repair after earthquake
Ease of field application
- epoxy, tape and fabric
The polyethylene fibers have the same property drawbacks as aramids, but also suffer from low melting temperature which limits
their use to composites that cure or operate below 149°C (300°F) and a susceptibility to degradation by ultraviolet light exposure.
4 lntvodmction, composite basics and road map
Both of these types of fibers have wide usage in personal protective armor. In spite of the drawbacks, production of both of these fibers is enjoying strong worldwide growth. Boron fibers (Chapter 8), the first advanced composite fibers to be used on production aircraft, are produced as individual monofilaments upon a tungsten or carbon substrate by pyrolytic reduction of boron trichloride (BC1,) in a sealed glass chamber. The relatively large cross section fiber is used today primarily in metal matrix composites which are processed at temperatures which would attack carbon/graphite fibers.
1.2 MATRIX SYSTEMS
If parallel and continuous fibers are combined with a suitable matrix and cured properly, unidirectional composite properties such as those shown on Table 1.3 are the result. The functions and requirements of the matrix are to: 0
0 0
0
0
keep the fibers in place in the structure; help to distribute or transfer loads; protect the filaments, both in the structure and before and during fabrication; control the electrical and chemical properties of the composite; carry interlaminar shear.
Table 1.3 Properties of typical unidirectional graphite/epoxy composites (Fiber volume fraction, V , = 0.62)
Elastic constants
High strength
High modulus
GPa (psi x I O 6 )
GPa (psi x IO6)
145 (21) 9.6 (1.4) 5.8 (0.85) 0.30
220 (32) 6.9 (1.0) 4.8 (0.7) 0.30
~~
Longitudinal modulus, E, Transverse modulus, E , Shear modulus, G , Poisson’s ratio (dimensionless)u ~ , ~~~
.~
~~
~
~~
~~
~~
Strength properties ~
Longitudinal tension, Ft”, Transverse tension, FtUT Longitudinal compression, FCUL Transverse compression, FCUT Inplane shear, PLT Interlaminar shear, F’,””
~~~
~~~~
~
MPa ( Z 0 3 psi)
MPa (lo3psi)
2139 (310) 54 (7.8) 1724 (250) 76 (11) 87 (12.6) 128 (18.5)
760 (110) 28 (4) 690 (100) 170 (25) 70 (10) 70 (10)
~
Ultimate strains -
%
-~
Longitudinal tension, Transverse tension, Longitudinal compression, ECUL Transverse compression, EC1lT Inplane shear
1.4 0.67 0.9 3.6 2.0
0.3 0.4 0.3 2.8
1600 (0.056)
1700 (0.058)
-0.079 (-0.044)
-0.54 (-0.3)
21.6 (12)
58 (32)
-
Physical properties Density, kg/m3 (Ib/in3) Longitudinal CTE, ye/K (pe/OF) Transverse CTE ~ E / (pe/OF) K
From References 1, 2 and 3; CTE = coefficient of thermal expansion
Matrix systems 5 The needs, or desired properties of the matrix, that depend on the purpose of the structure are: 0 0 0 0 0
0 0
0
0
0
0
minimize moisture absorption; have low shrinkage; Must wet and bond to fiber; low coefficient of thermal expansion; must flow to penetrate the fiber bundles completely and eliminate voids during the compacting/curing process; have reasonable strength, modulus and elongation (elongation should be greater than fiber); must be elastic to transfer load to fibers; have strength at elevated temperature (depending on application); have low temperature capability (depending on application); have excellent chemical resistance (depending on application); be easily processable into the final composite shape; have dimensional stability (maintain its shape).
There are many matrix choices available; each type has impact o n the processing technique, physical and mechanical properties and environmental resistance of the finished
composite. The common thermoset matrices for composites include the following: 0
0 0 0
polyester and vinylesters (Chapter 2); epoxy (Chapter 3); bismaleimide (BMI) (Chapter 4); polyimide (Chapter 4); cyanate ester and phenolic triazine (Chapter 5).
Each of the resin systems has some drawbacks, which must be accounted for in design and manufacturing plans. Polyester matrices have been in use for the longest period, and are used in the widest range and greatest number of structures. The usable polymers may contain up to 50% by weight of unsaturated monomers and solvents such as styrene. Polyesters cure via a catalyst (usually a peroxide) resulting in an exothermic reaction, which can be initiated at room temperature. The most widely used matrices for advanced composites have been the epoxy resins. These resins cost more than polyesters and do not have the high temperature capability of the bismaleimides or polyimides, but because of the advantages shown in Table 1.4 they are widely used.
Table 1.4 Selection criteria for epoxy resin systems
Advantages
Disadvantages
Adhesion to fibers and to resin
Resins and curatives somewhat toxic in uncured form
No by-products formed during cure
Absorb moisture Heat distortion point lowered by moisture absorption
Low shrinkage during cure High or low strength and flexibility Solvent and chemical resistance Resistance to creep and fatigue Solid or liquid resins in uncured state Wide range of curative options Adjustable curing rate Good electrical properties
Change in dimensions and physical properties due to moisture absorption Limited to about 200°C upper temperature use (dry) Difficult to combine toughness and high temperature resistance High thermal coefficient of expansion High degree of smoke liberation in a fire May be sensitive to ultraviolet light degradation Slow curing
6 Introduction, composite basics and road map There are two resin systems in common use for higher temperatures, bismaleimides and polyimides. New designs for aircraft demand a 177°C (350°F) operating temperature not met by the other common structural resin systems. The primary bismaleimide in use is based on the reaction product from methylene dianiline (MDA) and maleic anhydride: bis (4-maleimidophenyl) methane (MDA BMI). Two newer resin systems have been developed and have found applications in widely diverse areas. The cyanate ester resins, marketed by Ciba-Geigy, have shown superior dielectric properties and much lower moisture absorption than any other structural resin for composites. The dielectric properties have enabled their use as adhesives in multilayer microwave printed circuit boards, and the low moisture absorbance have caused them to be the resin of universal choice for structurallystable spacecraft components. The phenolic triazine (PT) resins also have superior elevated temperature properties, along with excellent properties at cryogenic temperatures. Their resistance to proton
ROVING
radiation under cryogenic conditions was a prime cause for their choice for use in the superconducting supercollider, subsequently canceled by the US Congress. Polyimides are the highest temperature polymer in general advanced composite use with a long term upper temperature limit of 232°C (450°F) or 316°C (600°F). Two general types are: condensation polyimides, that release water during the curing reaction, and addition type polyimides with somewhat easier process requirements. 1.3 FIBER MATRIX SYSTEMS
The end user sees a composite structure. Someone else, probably a prepregger, combined the fiber and the resin system and someone else caused the cure and compaction to result in a laminated structure. A schematic of the steps to arrive at a finished composite from the initial fiber is shown in Fig. 1.1. In many cases, the end user of the structure has fabricated the composite from prepreg, which is a low-temperature-stable combination
WEAUE?
N?
COLLIMRTE
UNI TRPE
A
Fig. 1.1 Manufacturing steps in composite structure.
Fiber matrix systems 7 of the resin, its curing agents and the fiber. The three types of continuous fibers, roving, tape and woven fabric available as prepregs give the end user many options in terms of design and manufacture of a composite structure. Although the use of dry fibers and impregnation at the work (i.e. filament winding, pultrusion or hand lay-up) is very advantageous in terms of costs; there are many advantages to the use of prepregs as shown in Table 1.5, particularly for the manufacture of modem composites.
The prepreg process for thermoset matrices can be accomplished by feeding the fiber continuous tape, woven fabric or roving through a resin-rich solvent solution and then removing the solvent by hot tower drying. The excess resin is removed via a doctor blade or metering rolls and then the product is staged to the cold-stable prepreg form, (B stage) (Fig. 1.2). The newer hot melt procedure for prepregs is gradually replacing the solvent method because of environmental concerns. A film of resin that has been cast hot onto release paper
Table 1.5 Advantages of prepregs over wet impregnation
Prepregs reduce the handling damage to dry fibers Improve laminate properties by better dispersion of short fibers Prepregs allow the use of hard-to-mix or proprietary resin systems Allow more consistency because there is a chance for inspection before use Heat curing provides more time for the proper laydown of fibers and for the resin to move and degas before cure Increased curing pressure reduces voids and improves fiber wetting Most prepregs have been optimized as individual systems to improve processing
Release Poly
Unwind
Prepreg Wind
Pump and Reservoir
Unwind
Fig. 1.2 Schematic of typical solvent prepregging process. (Adapted from Reference 2.)
8 Introduction, composite basics and road map
0
0
Pauer Paper
T
Doctor Plate 1 Impregnation Zone Creel
Paper
Plate 2
Take-up Prepreg Windup
Chill Plate
Fig. 1.3 Schematic of typical film impregnating process. (Adapted from Reference 2.)
is fed, along with the reinforcement, through a series of heaters and rollers to force the resin into the reinforcement. Two layers of resin are commonly used so that a resin film is on both sides of the reinforcement; one of the release papers is removed and the prepreg is then trimmed, rolled and frozen (Fig. I.3)2.The solvent technique has been largely replaced for advanced fibers because of environmental pollution concerns and a need to exert better control over the amount of resin on the fiber. 1.3.1 UNIDIRECTIONAL PLY PROPERTIES
The manufacturer of the prepreg reports an areal weight for the prepreg and a resin percentage, by weight. Each of the different fibers has a different density, resulting in a composite of different density at the same fiber volume percentage. Since fiber volume is used to relate the properties of the manufactured composites, the following equations can be used to convert between weight fraction and fiber volume.
where:
Wf = weight fraction of fiber wf = weight of fiber wc= weight of composite pf = density of fiber p, = density of composite uf = volume of fiber u, = volume of composite Vf = volume fraction of fiber V, = volume fraction of matrix p, = density of matrix. A percentage fiber that is easily achievable and repeatable in a composite and convenient for reporting mechanical and physical properties for several fibers is 60%. The properties of unidirectional fiber laminates are shown in Table 1.3 for carbon/graphite/epoxy. Values for the other fibers can be seen in their respective chapters. These values are for individual lamina or for a unidirectional composite, and they represent the theoretical maximum (for that fiber volume) for longitudinal in plane properties. Transverse, shear and compression properties will show maxima at different fiber volumes and for different fibers, depending on how the matrix and fiber interact. These values can be used to calculate the properties of a laminate which has fibers oriented in several directions. To do that, the methods of description for ply orientation must be introduced.
Quasi-isotropic laminate 9 1.4 PLY ORIENTATIONS, SYMMETRY AND
BALANCE
lined to indicate that half of it lies on either side of the plane of symmetry (Fig. 1.4(f)).
1.4.1 PLY ORIENTATIONS
One of the advantages of using a modern composite is the potential to orient the fibers to respond the load requirements. This means that the composite designer must show the material, the fiber orientations in each ply, and how the plies are arranged (ply stackup). A 'shorthand' code for ply fiber orientations has been adapted for use in layouts and studies. Each ply (lamina)is shown by a number representing the direction of the fibers in degrees, with respect to a reference ( x ) axis. 0" fibers of both tape and fabric are normally aligned with the largest axial load (axis) (Fig. 1.4(a)). Individual adjacent plies are separated by a slash in the code if their angles are different (Fig. 1.4@)). The plies are listed in sequence, from one laminate face to the other, starting with the ply first on the tool and indicated by the code arrow with brackets indicating the beginning and end of the code. Adjacent plies of the same angle of orientation are shown by a numerical subscript (Fig. 1.4(c)). When tape plies are oriented at angles equal in magnitude but opposite in sign, (+) and (-) are used. Each (+) or (-) sign represents one ply. A numerical subscript is used only when there are repeating angles of the same sign. Positive and negative angles should be consistent with the coordinate system chosen. An orientation shown as positive in one right handed coordinate system may be negative in another. If the y and z axis directions are reversed, the f 45 plies are reversed (Fig. 1.4(d)). Symmetric laminates with an even number of plies are listed in sequence, stating at one face and stopping at the midpoint. A subscript 'S' following the bracket indicates only one half of the code is shown (Fig. 1.4(e)). Symmetric laminates with an odd number of plies are coded as a symmetric laminate except that the center ply, listed last, is over-
1.4.2 SYMMETRY
The geometric midplane is the reference surface for determining if a laminate is symmetrical. In general, to reduce out-ofplane strains, coupled bending and stretching of the laminate and complexity of analysis, symmetric laminates should be used. However, some composite structures (e.g. filament wound pressure vessels) can achieve geometric symmetry so that symmetry through a single laminate wall is not necessary, if it constrains manufacture. To construct a midplane symmetric laminate, for each layer above the midplane there must exist an identical layer (same thickness, material properties, and angular orientation) below the midplane (Fig. 1.4(e)). 1.4.3 BALANCE
All laminates should be balanced to achieve inplane orthotropic behavior. To achieve balance, for every layer centered at some positive angle +e there must exist an identical layer oriented at -8 with the same thickness and material properties. If the laminate contains only 0" and/or 90" layers it satisfies the requirements for balance. Laminates may be midplane s p metic but not balanced and vice versa. Figure 1.4(e) is symmetric and balanced whereas Fig. 1.4(g)is balanced but unsymmetric . 1.5 QUASI-ISOTROPICLAMINATE
The goal of composite design is to achieve the lightest, most efficient structure by aligning most of the fibers in the direction of the load. Many times there is a need, however, to produce a composite which has some isotropic properties, similar to metal, because of multiple or undefined load paths. A 'quasi-isotropic' laminate lay-up accomplishes this for the x and y planes only; the z or through-the-laminate-
-
10 Introduction, composite basics and road map 90"
Reference Axis
lz;
90"
Tool side
,/
.-,
Tape Laminate
0"
1
I I P
I
45'
90' -45'
-450
w
\
\
0"
+450
[0/903/0]
P
I
90" +45" -45" -45" +45" 90" V
0"
[0/9O]s Typical Callout
T
[0/90/*45]s
Typical Callout
Line of Symmetry
I
Tape and Fabric Laminate [ 0/f45/To1 s. Typical Callout
Line of Symmetry
Fig. 1.4 Ply orientations, symmetry and balance. (Continued on next page)
0
0"
L
Methods of analysis 11 Tape Laminate
p, +45" -45"
[0/90/f45/i452/9 0/ 01 Typical Callout
-45"
+45"
Fabric Laminate
I
0".90" I
j
0",90"
I
[(0,90)/(~45)/(0,90)] Typical Callout
h)
Fig. 1.4 Ply orientations, symmetry and balance. (Continued)
thickness plane is quite different and lower. 1. arrive at quick values to determine if a comMost laminates produced for aircraft applicaposite is feasible; tions have been, with few exceptions, 2. arrive at values for insertion into computer 'quasi-isotropic'. As designers become more programs for laminate analysis or finite eleconfident and have access to a greater database ment analysis; with fiber-based structures, more applications 3. check on the results of computer analysis. will evolve. For a quasi-isotropic laminate, the The rule of mixtures holds for composites. The following are requirements: micromechanics formula to arrive at the 0 It must have three layers or more. Young's modulus for a given composite is: 0 Individual layers must have identical stiffEc = V,E, + Vm Em ness matrices and thicknesses. 0 The layers must be oriented at equal angles. and v,+ vm= 1 For example, if the total number of layers is = V ,E , + Em (1- V,) (1.3) M , the angle between two adjacent layers should be 360"ln. If a laminate is con- where structed from identical sets of three or more Ec = composite or ply Young's modulus in layers each, the condition on orientation tension for fibers oriented in direction of must be satisfied by the layers in each set, applied load for example: ( O o / + 60"), or ( O o / + 45"/90)s. V = volume fraction of fiber ( f ) or matrix (m) E = Young's modulus of fiber ( f ) or matrix 1.6 METHODS OF ANALYSIS (m). There are a number of methods in common But, since the fiber has much higher use for the analysis of composite laminates. Young's modulus than the matrix, the second The use of micromechanics, i.e. the application part of the equation can be ignored. of the properties of the constituents to arrive at E, >> Em the properties of the composite ply can be used to: Ec = E,V, (1.4)
12 Introduction, composite basics and road map appropriate for a particular application. Figure 1.5 shows the progression of physical properties for Young’s modulus in tension, E, (fiber), E, (lamina) and Ex,, (laminate), longitudinal tensile strength, and coefficient of thermal expansion a where the subscripts L and X stand for in-plane in the principal fiber direction and t and Y stand for the transverse direction for a theoretical high strength (from Ec = (3/8) E,V, (1.5) Table 1.3) carbon/graphite fiber composite The quasi-isotropic modulus, E, of a composite from the fiber to the laminate. The values laminate is (3/8)E,+(5/8)EZ where E,, is the decrease or are ’translated’in a logical fashion modulus of the lamina in the fiber direction and and reflect the law of mixtures. The analysis is E, is the transverse modulus of the lamina3. relatively simple for modulus dominated The transverse modulus for polymeric-based properties but strength-dominated values composites is a small fraction of the longitudinal must be treated in light of one of several failmodulus (see E, in Table 1.3)and can be ignored, ure theories and changes in the thermal for preliminary estimates, resulting in a slightly coefficient of expansion are not predictable lower-than-theoretical value for Ec for a quasi- from laws of mixtures. Other factors which isotropic laminate. This approximate value for enter into the translation efficiency are: comthe quasi-isotropic modulus represents the patibility of the resin system with the fiber and lower limit of composite modulus. It is useful in the fiber finish, strain-to-failure of the resin comparing of composite properties to those of system and the damage the fiber undergoes metals and in establishing if a composite is during impregnation, laydown and cure.
This is the basic rule of mixture and represents the highest Young’s modulus composite, where all fibers are aligned in the direction of load. The minimum Young’s modulus for a reasonable design (other than a preponderance of fibers being orientated transverse to the load direction) is the quasi-isotropic composite and can be approximated by:
.6 GPa, FT‘“ =54 MPa
E x = 76 GPa a x = 4.98peK
r
E y = 76 GPa a y = 4.98~ E K
a2
>ay
Fig. 1.5 The anatomy of a composite laminate.
Composite fabrication techniques Table 1.6 High-strength carbon/graphite laminate
properties Laminate
(0/90,/0) (90/0,/90) (02/902/OJ (0,/~45,/0,) (0/+45/90)>
Aluminum
Longitudinal modulus E,, (GPa) 76.5 76.5 98.5 81.3 55.0 41.34
Bending modulus, E , (GPa) 126.8 26.3 137.8 127.5 89.6 41.34
Shear modulus, G,, (GPa) 5.24 5.24 5.24 21.0 21.0 27.56
Table 1.6 shows mechanical values for several composite laminates with the fiber of Table 1.3 and a typical resin system. The first and second entries are for simple 0/90 laminates and show the effect of changing the position of the plies. The effect of increasing the number of 0 plies is shown next and the final two laminates demonstrate the effect of +45 plies on mechanical properties, particularly the shear modulus. The last entry is a quasi-isotropic laminate. These laminates are then compared to a typical aluminum alloy. When employing the data extracted from tables, some caution should be observed by the reader. The values seen in many tables of data may not always be consistent for the same materials or the same group of materials from several sources for the following reasons: 1. Manufacturers have been refining their production processes so that newer fibers may have greater strength or stiffness. These new data may not be reflected in the compiled data. 2. The manufacturer may not be able to change the value quoted for the fiber because of government or commercial restrictions imposed by the specification process of his customers. 3. There are many different high-strength fibers commercially available. Each manufacturer has optimized their process to maximize their mechanical properties and each process may differ from that of the
13
competitor, so vendor values in a generic class may differ widely. 4. Most tables of values are presented as 'typical values'. Those values and the values that are part of the menu of many computer analysis programs should be used with care. Each user must find their own set of values for design, develop useful design allowables, and apply appropriate 'knock down' factors, based on the operating environments expected in service. (Chapter 33 and Appendix A give guidelines.) 1.7 COMPOSITE FABRICATION TECHNIQUES
The goals of the composite manufacturing process are to: 0
achieve a consistent product by controlling fiber thickness; - fiber volume; - fiber directions; minimize voids; reduce internal residual stresses; process in the least costly manner. -
0 0 0
The procedures to reach these goals involve iterative processes to select the three key components: 0 0 0
composite material and its configuration; tooling; process.
Once material selection has been completed, the first step leading to the acceptable composite structure is the selection of tooling, which is intimately tied to process and material. For all curing techniques the tool must be: 0
0
0
0
strong and stiff enough to resist the pressure exerted during cure; dimensionally stable through repeated heating and cooling cycles; light enough to respond reasonably quickly to the changes in cure cycle temperature and to be moved in the shop; leakproof so that the vacuum and pressure cycles are consistent.
14 Introduction, composite basics and road map
The tool face is commonly the surface imparted to the outer surface of the composite and must be smooth, particularly for aerodynamic surfaces. The other surface frequently may be of lower finish quality and is imparted by the disposable or reusable vacuum bag. This surface can be improved by the use of a supplemental metal tool known as a caul plate. (Press curing, resin transfer molding, injection molding and pultrusion require a fully closed or two sided mold). Figure I. 6 shows the basic components of the tooling for vacuum bag or autoclave processed components and Table 1.7 shows the function of each part of the system. Tooling options have been augmented by 3
2
12
13
4
14
the introduction of elastomeric tooling wherein the thermal expansion of an elastomer provides some or all of the pressure curing cure, or a rubber blanket is used as a reusable vacuum bag. The volumetric expansion of an elastomer can be used to fill a cavity between the uncured composite and an outer mold. The use of elastomeric tooling can provide the means for fabricating complex box-like structures such as integrally stiffened skins with a co-cured substructure in a single curing operation. Tooling (Chapter 25) and the configuration of the reinforcement have a great influence on the curing process selected and vice-versa. The 5
6
9
8
7
10
9
11
Fig. 1.6 Typical vacuum bag lay-up components. Table 1.7 Functions of vacuum bag components
Component *
Functions
-
1 2 3 4 5 6 7 8 9 10 11 12
13 14
Bag sealant Vacuum fitting and hardware Bagging film Open weave breather mat Polyester tape (wide) Polyester tape (narrow) Caul sheet Perforated release film Non-perforated release film Peel ply Laminate 1581-styleglass breather manifold 1581 style glass bleeder ply Stacked silicone edge dam
* numbers refer to Fig. 1.6
Temporarily bonds vacuum bag to tool Exhausts air, provides convenient connection to vacuum pump Encloses part, allows for vacuum and pressure Allows air or vacuum transfer to all of part Holds other components of bag in place Holds components in place Imparts desired contour and surface finish to composite Allows flow of resin or air without adhesion Prevents adhesion of laminate resin to tool surface Imparts a bondable surface to cured laminate Allows transfer of air or vacuum Soaks up excess resin Forces excess resin to flow vertically, increasing fluid pressure
Composite fabrication techniques 15 probable reinforcement configuration that facilitates the completion of the finished composite is shown on Table 1.8. The choice between unidirectional tape and woven fabric has frequently been made on the basis of the greater strength and modulus attainable with the tape particularly in applications which compression strength is important. There are other factors that should be included in the trade, as shown in Table 1.9. 1.7.1 LAY-UP TECHNIQUE
Lay-up techniques along with composite cure control have received the greatest attention for processing. In efforts to reduce labor costs of composite fabrication, to which lay-up (Chapter 16) has traditionally been the largest contributor, mechanically assisted, controlled tape laying and automated integrated manu-
Table 1.8 Common reinforcement configuration for the manufacturing process
Reinforcement Prepreg Prepreg Prepreg Other, configuration tape or (dry) or (dry) woven tow woven preforms, or non- chopped woven fibers fabric Handlay-up Automatic tape laydown
X X
x, (XI
x
Filament winding
x, (X)
xm
xm
Resin transfer molding
(XI
(X)
X
Pultrusion
(X)
Fiber placement
X
X
X
Table 1.9 Fabric compared with tape reinforcement
Tape advantages
_
_
_
_
_
~
-
~
Tape disadvantages
Best modulus and strength efficiency
Poor drape on complex shapes
High fiber volume achievable
Cured composite more difficult to machine
Low scrap rate
Lower impact resistance
No discontinuities
Multiple plies required for balance and symmetry
Automated lay-up possible
Higher labor costs for hand lay-up
Available in thin plies Lowest cost prepreg form Less tendency to trap volatiles
Fabric advantages
Fabric disadvantages
~
_ _ ~ _ _ _ - - ~ _ _
Better drape for complex shapes
Fiber discontinuities (splices)
Single ply is balanced and may be essentially symmetric
Less strength and modulus efficient
Can be laid up without resin
Lower fiber volume than tape
Plys stay in line better during cure
More costly than tape
Cured parts easier to machine
Greater scrap rates
Better impact resistance
Warp and fill properties differ
Many forms available
Fabric distortion can cause part warping
16 Introduction, composite basics and road m a p facturing systems have been developed. Table Generally, the percent matrix weight is higher 1.10 shows some of the considerations for before cure initiation; the matrix flows out of choosing a lay-up technique. the laminate and takes the excess resin with In addition to any cost savings by the use of the potential voids. An arbitrary 1%void limit an automated technique for long production has been adopted for most autoclaved comruns, there are two key quality assurance fac- posites; filament wound and pultruded tors which validate the automated techniques. composites will have higher void volumes They are: greatly reduced chance that release depending upon the application. An autoclave is essentially a closed, prespaper or film could be retained, which would destroy shear and compressive strength if surized oven; many common epoxy laminates undetected, and reduced probability of the are cured at an upper temperature of 177°C addition or loss of an angle ply which would (350°F) and 6 MPa (100 psi). Autoclaves are cause warping due to the laminate’s lack of still the primary tool in advanced composite symmetry and balance. processing and have been built up to 16 m (55 All curing techniques use heat and pressure feet) long at 6.1 m (20 feet) diameter. Since to cause the matrix to flow and wet out all the autoclaves are expensive to build and operate, fibers before the matrix solidifies (Chapter 26). many other methods of curing, compacting Table 1.10 Considerations in composite lay-up technique
Considerat ion
Manual
Flat tape
Contoured tape
Orientation accuracy
Least accurate
Automatic
Somewhat dependent on tape accuracy and computer program
Ply count
Dependent on operator, count Mylars
Dependent on operator
Program records
Release film retention
Up to operator
Automatic
Automatic removal
Labor costs
High
86% improvement quoted
Additional improvement
Machine costs
N/A
Some costs
Approximately 1M$ or greater
Production rate
Low (1.5 Ib/h)
10 lb/h
Approximately same as flat tape
Machine ’up’ time
N/A
Not a consideration
Complex program and machine make this a consideration
Varying tape widths
Not a concern
Easily changed
Difficulty in changing
Tape lengths
Longer tapes more difficult
Longer is more economical
Longer tape is more economical
Cutting waste
Scrap on cutting
Less scrap
Least scrap due to back and forth laydown
Compaction pressure
No pressure
Less voids
Least voids
Programming
N/A
N/A
Necessary
Compositefabrication techniques 17 composites have been developed. The two newest and most attractive methods are fiber placement and resin transfer molding. 1.7.2 RESIN TRANSFER MOLDING
Previous discussions have centered on moving resin out of the laminate to reduce voids. Resin transfer involves the placement of dry fiber reinforcement into a closed mold and then injecting a catalyzed resin into the mold to encapsulate the reinforcement and form a composite (Chapter 20). The impetus for the use of this process comes from the large cost reductions that can be realized in raw materials and lay-up. The process can utilize low injection pressures i.e. 55 MPa (80 psi), therefore, the tooling can be lower cost plastic or a vacuum bag rather than metal.
a wind eye at speeds synchronized with the mandrel rotation, control winding angle of the reinforcement and the fiber lay-down rate. The reinforcement may be wrapped in adjacent bands or in repeating bands that are stepped the width of the band and that eventually cover the mandrel surface. Local reinforcement can be added to the structure using circumferential windings, local helical bands, or by the use of woven or unidirectional cloth. The wrap angle can be varied from low angle helical to high angle circumferential or 'hoop', which allows winding from about 4"-90" relative to the mandrel axis; newer machines can 'place' fiber at 0". 1.7.4 FIBER PLACEMENT
Fiber placement, initially developed by Hercules Aerospace Co., is a cross between filament winding and automatic tape laydown, 1.7.3 FILAMENT WINDING retaining many of the advantages of both. The Filament winding is a process by which con- natural outgrowth of adding multiple axes of tinuous reinforcements in the form of rovings control to filament winding machines results or tows (gathered, untwisted strands of fiber) in control of the fiber laydown so that non axiare wound over a rotating mandrel. The man- symmetric surfaces can be wound. This drel can be cylindrical, round or any other involves the addition of a modified tape layshape as long as it does not have re-entrant down head to the filament winding machine curvature. Special machines (Fig. and much more. The Cincinnati-Milacron - 1.7) traversing machine additions include in-process compaction, individual tow cut/start capabilities, a resin tack control system, differential tow payout, low tension on fiber and enhanced offf 1 iine programming (Chapter 22).
I/ 11
1.7.5 PULTRUSION
Pultrusion is an automated process for the I manufacture of constant volume/shape profiles from composite materials (Chapter 23). The composite reinforcements are continuously pulled through a heated die and shaped and cured simultaneously. If the cross-sectional shape is conducive to the process, it is p f r 4 the fastest and most economical method of Fig. 1.7 The helical filament wound ply. (Courtesy composite production. Straight and cured conof Westinghouse Electric Co., Marine Division.) figurations can be fabricated with square,
cF
I
18 Introduction, composite basics and road map round, hat-shaped, angled 'I' or 'T'-shaped cross-sections from vinylester, polyester, or epoxy matrices with E and S-glass, Kevlar and graphite reinforcements.. The curing is effected by combinations of dielectric preheating and microwave or induction (with conductive reinforcements like carbon graphite) while the shape traverses the die. 1.7.6 BRAIDING, WEAVING AND OTHER PREFORM TECHNIQUES
I
3
Fig. 1.8 The unidirectional ply.
Braiding, weaving, knitting and stitching represent methods of forming a shape, generally be the same in any transverse direction. This is referred to as preforming, with the composite the transverse isotropy assumption; it is fibers before impregnation (Chapter 18). The approximately satisfied for most unidirecshape may be the final product or some inter- tional composite plies. These properties are typically modified by mediate form such as a woven fabric. The braiding process is continuous and is transformation relative to the laminate axis amenable to round or rectangular shapes or where these may not be the same as the ply smooth curved surfaces and can transition axes. In a multidirectional laminate there can be easily from one shape to another. The other fabric preforming techniques are as many as 21 stiffness constants. Strength preweaving, knitting and the non-structural dictions are equally as complicated because of stitching of unidirectional tapes. Stitching sim- directional differences, i.e. compression is not ply uses a non-structural thread, such as nylon always equal to tension, and because the sevor Dacron, to hold dry tapes at selected fiber eral failure theories are complex. As the angles. Preforming in this manner results in a complexity of the matrix calculations increase, higher-cost raw material but saves labor costs it becomes evident that errorless mathematical for orientation of individual lamina. The manipulations are impossible without the aid stitched preform has known, stable fiber ori- of computers. Chapters 30 and 32 elaborate on entations similar to woven fabric, without the the techniques of laminate analysis and the crossovers which could reduce compressive applications of laminates to structures strength. 1.9 DESIGN OF COMPOSITES 1.8 MECHANICS OF COMPOSITE MATERIALS
The 1,2,3 axes in Fig. 1.8 are special and are called the ply axes, or material axes. The 1 axis is in the direction of the fibers, and is called the longitudinal axis or the fiber axis. The longitudinal axis is typically the highest stiffness and strength direction. Any direction perpendicular to the fibers (in the 2,3 plane) is called a transverse direction. Sometimes, to simplify analysis and test requirements, ply properties are assumed to
The design process for composites involves both laminate design and component design and must also include considerations of manufacturing process and eventual environmental exposure. These steps are all interdependent with composites and the most efficient design must involve true concurrent engineering. Figure 1.9 shows the various concerns that should be a part of the composite design process at the initiation of the design process, and continuously from there on.
Design of composites 19 1.9.1 LAMINATE DESIGN RECOMMENDATIONS
1. Take advantage of the orthotropic nature of the fiber composite ply. 0 To carry in-plane tensile or compressive loads align the fibers in the directions of these loads. 0 For in-plane shear loads, align most fibers at -c 45" to these shear loads. 0 For combined normal and shear in-plane loading provide multiple or intermediate ply angles for a combined load capability. 2. Intersperse the ply orientations. 0 If a design requires a laminate with 16 plies at *45", 16 plies at 0", and 16 plies at 90°, use the interspersed design (90,/ -c 45,/0,),s rather than (90,/ .+ 45,/10,)s. Concentrating plies at nearly the same angle (0" and 90" in the above example) provides the opportunity for large matrix cracks to form. These produce lower laminate allowables, probably because large cracks are more injurious to the fibers, and more readily form delaminations than the finer cracks occurring in interspersed laminates. 0 If a design requires all 0" plies, some 90" plies (and perhaps some off-angle plies ) should be interspersed in the laminate to provide some biaxial strength and stability and to accommodate unplanned Composite Material Environmental Considerations
Component
Fig. 1.9 Design considerations for composites.
0
loads. This improves handling characteristics, and serves to prevent large matrix cracks from forming. Locally reinforce with fabric or mat in areas of concentrated loading. (This technique is used to locally reinforce pressure vessel domes). Use fabric, particularly fiberglass or Kevlar, as a surface ply to restrict surface (handling) damage. Ensure that the laminate has sufficient fiber orientations to avoid dependence on the matrix for stability. A minimum coverage of 6 to 10% of total thickness in 0, ?45", 90" directions is recommended.
3. Select the lay-up to avoid mismatch of properties of the laminate with those of the adjoining structures - or provide a shear/separator ply. Poisson's ratio: if the transverse strain of a laminate greatly differs from that of adjoining structure, large interlaminar stresses are produced under load. Coefficient of thermal expansion: temperature change can produce large interlaminar stresses if coefficient of thermal expansion of the laminate differs greatly from that of adjoining structure. 0 The ply layer adjacent to most bonded joints should not be perpendicular to the direction of loading. Thicken the composite in the joint area, soften the composite by adding fiberglass or angle plies and select the highest strain-capability adhesive. 4. Use multiple ply angles. Typical composite laminates are constructed from multiple unidirectional or fabric layers which are positioned at angular orientations in a specified stacking sequence. From many choices, experience suggests a rather narrow range of practical construction from which the final laminate configuration is usually selected. The multiple layers are usually oriented in at least two different angles, and possibly three or four; (go, O0/&", or
20 Introduction, composite basics and road map
attempt to standardize the raw materials and their test methods by publication of specifications (Appendix A). However, these standards have not reached the level of use to allow complete dependence upon them without supplier-user interaction and user testing. The fabricators of composites will rely on specifications for control of fiber, resin and/or the prepreg. Many prepreg resin and fiber Further suggestions can be seen in Chapter 31. vendors will certify only to their own specifications which may differ from those shown; users should consult the vendors to determine 1.10 COMPOSITE TESTING what certification limits exist before commitTo ensure consistent, reproducible compo- ting to specification control. nents, three levels of testing are employed: As part of raw materials verification, comincoming materials testing, in-process testing posite design effort and final product and control and final structure verification. verification mechanical testing of composite test specimens will be performed. The testing of composite materials offers unique chal1.10.1 INCOMING MATERIALS TESTING lenges because of the special characteristics of Incoming materials testing seeks to verify the composites. Factors not considered important conformance of the raw materials to specifica- in metals testing are very important in testing tions and to insure processibility. The levels of composites (Chapters 34,39). knowledge of composite raw materials do not approach those for metals, which can be bought to several consensus specifications and REFERENCES will appear generally identical although pur- 1. Foral, R.F. and Peters, S.T., Composite chased from many manufacturers. Although Structures and Technology Seminar Notes, 1989 there are fewer suppliers for composite raw 2. Hercules Data Sheet for AS-4/3901-6 prepreg H050-377/GF Prod Hdbk (4)/jc/2 materials, the numbers of permutations of resins, fibers and manufacturers prevents the 3. Agarwal, B.D. and Broutman, L.J., Analysis and Performance of Fiber Composites 2nd edn, John kind of standardization necessary to be able to Wiley and Sons, New York, 1990 p. 103 buy composite raw materials as if they were 4. Mayorga, G.D. in International Encyclopedia of alloys. ASTM (American Society for Testing Composites, (ed. S.M. Lee) Vol 4, VCH and Materials), SAE/AMS/NOMETCOM Publishers, N.Y., N.Y., 1991 (Society of Automotive Engineers, Aeronautical 5 . Tsai, S.W. and Pagano, N.J. in Composite Materials Workshop, (eds. S.W. Tsai, J.C. Halpin Materials Standards/ Nonmetallic Materials and N.J. Pagano), Technomic Publishing Co., Committee) and SACMA (Suppliers of Lancaster, PA, 1978, p. 249 Advanced Composite Materials Association) 0 ° / ~ 0 / 9 0 cover 0 most applications, with 0 between 30 and 60 degrees). Unidirectional laminates are rarely used except when the basic composite material is only mildly orthotropic (e.g. certain metal matrix applications) or when the load path is absolutely known or carefully oriented parallel to the reinforcement (e.g. stiffener caps).
2
POLYESTER AND VINYL ESTER RESINS Frank A. Cassis and Robert C. Talbot
2.1 INTRODUCTION AND HISTORY
Organic polymers are divided into two types, reinforced-thermoplastic and thermoset. With thermoset polymers such as unsaturated polyesters and vinyl esters, a chemical reaction cross links the material so that it cannot be returned to liquid form. Other common thermosetting polymers include epoxy and phenolic resins. Thermoset plastics made with polyester and vinyl ester resins represent the major portion of the reinforced plastic composites industry today. Early workers on unsaturated polyesters soon learned that despite the possession of reactive double bonds, these resins were sluggish in reacting with themselves. Even with effective catalysts, they still required high temperatures and lengthy cure times to complete the cross linking reaction. The key to modern day application of unsaturated polyesters was the discovery by Carlton Ellis in 1937l that the addition of reactive monomers, such as styrene, gave mixtures that would copolymerize many times faster than homopolymerization. The styrene addition produced the added benefit of an easily handled liquid material that could be pumped, transported and fabricated into a finished plastic by a myriad of molding processes. Developments during the 1940s accelerated the commercial applicability of unsaturated polyesters to the position they hold today. Styrene became readily available and lower in cost as a result of the US Government's sponHandbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
sored production of styrene-butadiene rubber. At the same time, scientists found that styrenated polyesters could yield high strength, light weight structures when reinforced with glass fibers. They also learned that fiberglassreinforced polyesters had excellent electrical properties and that large structures could be molded at low pressures with low cost tooling. As a result, commercial development proceeded rapidly after World War I1 with materials and molding research moving in many directions. In the 20 years that followed, polyester and vinyl ester resins matured rapidly and by the mid-l970s, the composites fabricator and end user had numerous options with these matrix systems to achieve the desired properties in the finished part. 2.2 POLYESTER RESINS
The reaction of an organic acid with an alcohol results in the formation of an ester. By using a difunctional acid and a difunctional alcohol (glycol)a linear polyester is produced (Fig. 2.1). 0 0 II II H-(-0 - C - R - C - 0
- R -)" -OH
Fig. 2.1
Properties of polyesters can be varied by using different combinations of diacids and glycols. These products are thermoplastic polyesters and they are made with various acids and
Polyester resins 35 glycols such as the following: -
Acids Phthalic anhydride Isophthalic acid Terephthalic acid Adipic acid
GZycols Ethylene glycol Propylene glycol Neopentyl glycol Diethylene glycol
In the esterification reaction with maleic anhvdride, the unsaturated acid isomerizes to the fumarate structure which copolymerizes with styrene much faster than the maleate form. A high degree of isomerization to the fumarate structure is essential to produce an unsaturated polyester with high reactivity. Although the isomerization of maleic anhydride is usually from 65-95% in the esterification reaction, some commercial resins are deliberately formulated with the more expensive fumaric acid to obtain maximum reactivity with the monomer employed.
The reaction product of terephthalic acid and ethylene glycol is the well known polyethylene terephthalate (PET) which is used to make polyester fibers and polyester plastics such as clear plastic bottles for soft drinks. Unsaturated polyesters are produced by replacing part of the saturated diacid with an 2.2.1 UNSATURATED POLYESTER CLASSES unsaturated diacid such as maleic anhydride or fumaric acid (Fig. 2.2). The former is vastly Unsaturated polyesters are divided into types preferred since it is lower in cost, easily han- or classes depending on the structure of the dled and produces only half the water that basic building block. These are orthophthalic, would be generated in the reaction when isophthalic, terephthalic, bisphenol-fumarate, fumaric acid is used. chlorendic and dicyclopentadiene.
CH = CH
t
\
0 = c-0-c = 0 Maleic anhydride H
H II
I1
HOOC- C = C
- COOH
Fumaric acid
Fig. 2.2
The resultant polyester contains reactive double bonds (unsaturation) along the entire polyester chain, which becomes the site for the eventual cross linking to produce the cured plastic (Fig. 2.3). 0
0
0
HO
It II II I II HO ( C-R-C-O-R'-O-C-C=C-C-O-R'-O)n H I
H
Fig. 2.3
Orthophthalic resins These are commonly referred to as ortho or general purpose resins and are usually based on phthalic anhydride, maleic anhydride and propylene glycol. Since the acid groups in phthalic anhydride are on adjacent carbons of the benzene ring, it is very difficult to produce resin molecular weights as high as those achievable with isophthalic and terephthalic acid. Accordingly, resins made from phthalic anhydride have poorer thermal stability and chemical resistance than their iso/tere counterparts.
Isophthalic resins These resins are produced from isophthalic acid and are characterized by greater strength, heat resistance, toughness and flexibility than their ortho cousins. In isophthalic acid, the acid groups are separated by one carbon of the benzene ring which increases the opportunity to produce polymers with greater linearity and higher molecular weight in the esterification reaction (Fig. 2.4).
36 Polyester and viny ester resins
Phthalic anhydride
ester can be three times longer than for an is0 resin. As a result of this, researchers have turned to polyethylene terephthalate scrap from the previously mentioned fiber and plastic operations to develop an economical source of terephthalic polyesters. This scrap can be effectively depolymerized by using different amounts of propylene glycol at elevated temperatures. The glycolyzed product is then reacted with maleic anhydride and diluted with styrene monomer to produce a cost effective terephthalic polyester.
lsophthalic acid
0 c=o ‘OH
Bisphenol A fumarate resins
Terephthalic acid
Fig. 2.4
Therephthalic resins Unsaturated polyesters can be produced from terephthalic acid with the expectation that the resin property improvement obtained in going from phthalic anhydride to isophthalic acid will be matched in going from isophthalic to terephthalic acid. This, however, is not the case and terephthalic resins appear to offer only a slight advantage in heat distortion temperature over their isophthalic counterparts. Other important resin properties such as modulus, hardness and overall chemical resistance favor the is0 resins. Because of its lower solubility and poorer reactivity, therephthalic acid requires the use of esterification catalysts or pressure processing to produce a resin economically. Without these, processing time for a terephthalic poly-
H HO - C I - CH2 - O
o
l
@
These resins are unsaturated rigid polyesters made by reacting bisphenol A with propylene oxide to produce the glycol shown in Fig. 2.5. This propoxylated bisphenol A is then reacted with fumaric acid to form an unsaturated polyester. The bisphenol structure illustrated above imparts a high degree of hardness, rigidity and thermal stability to this particular resin. Chlorendic resins These unique polyester resins are based on HET acid (hexachlorocyclopentadiene) or the anhydride shown in Fig. 2.6. When reacted with an unsaturated acid and a stable glycol such as neopentyl, an extremely rigid unsaturated polyester results with outstanding thermal stability and resistance to oxidizing environments. The inherent chlorine in the resin chain imparts some fire retardancy as well.
H .. 0 - CH2 - C I - OH
I
I
CH3
CH3
Fig. 2.5
CI CQ ;?!
CI-c-CI
Fig. 2.6
-c=o
37
Hydrocarbon Solvents Table 2.72: Hexene-1 (4) cn2 = cn-cn2-cn2-cn2-cn,
FORMULA RESEARCH GRADE
PROPERTIES
PURE GRADE
TECHNICAL GRADE
'Literature values.
Table 2.73: cis-Hexene-2 (4)
FORMULA
cn3-c
Composition. weight percent ~
-
~
_
~
~
0.1 0.2 99.6
~
.
Purity by freezing point, mol % Freezing point, F Boiling point, F _ _ _ -~ -~ __Distillation range, F Initial boiling point Dry point 0 7 6 m Specific gravity of liquid at 6 at 2014 C API gravity at 60 F Density of liquid at 60 F. m a l Vapor pressure at 70 F. psia 100 F, psia -___ 130 F, psia index. 20/0 _ _ Refractive _ ._ Color. Saybolt Acidity, distillation residue Nonvolatile matter. gramdl00 ml -__Flash point. appoximate. F *Literature values.
FORMULA
0.1
-~
99.28 -222.04' 156.00'
0.6 920* _________ 0.68720' 5.760' 2.4' 4.9' 9.1' 1.39761. +30
1
-
cH3-cH cn-cH2-cn2-cn3 PURE GRADE
PROPERTIES
~
Hexene-1 ___-_-.___~_-tranr-Hexene-2 cis-Hexene.2 Hexe!Er-?p.Normal Hexane lsoolefins _ _ _ . _ __Heptene-l tranr-Heptene-3 cis.Heptene.3 tranr-Heptene.2 cis.Heptene2
_ _ _ _ _ _ ~
- c-cn2-cn2-cn, RESEARCH GRADE
PROPERTIES
~~
Table 2.74: Mixed 2-Hexenes (4)
~-
Composition, weight percent .. Hexenel . tiant-Hexene-2 cis Hexene-2 Hexener-3 Normal Hexane lsoolefins -_____ Heptene-l
Purity by freezing point. mol % Freezing point. F Boiling point, F Distillation range. F Initial boiling point Dry point Specific gravity of liquid at 60160 F at2014 c API gravity at 60 F Density of liquid a t 60 F. Ibdgal Vapor pressure a t 70 F. psis 100 F, psia 130 F, psia Refractive index, 2010 Color, Saybolt Acidity. distillation residue Nonvolatile matter, gramdl00 ml Flash point, approximate, F
-
-
trxe
0.8
21 ~ -
-.
_
.
-
~ -- - - -
-
-
03
.-
i:it99Omin
__
tranr-Heptene-3 cis-Heptene3 tranr-Heptene 2 ca-Heptene2
TECHNICAL GRADE
--
-
--__-
-
..
_
..
-
-
-
~
.~
.
~-
~~~
~~
~
____
__-.
155.0 155.1 0.686
155.0 __.___.__ 155.1 -. 0.684 ~
- . . ~
.
!.2._.
1.396 +30 _ -~ neutral _. 0.0005 -5
. ~
~ _ _ _ - . ~ -
75.4 5.69 2.4 ____ 5.0 -~ ... . -~ .~~
~
~~-
74.8 5.71 2.4
~-
5.0
- -~ .-
~~
.. ..
9.2 1.396 +30 neutral -. 0.0005 -5
-..- ~-.
__
.
-
~
__ ~
38 Polyester and viny ester resins In addition to tailoring the resin for specific applications by varying the building blocks, the properties of unsaturated polyesters can often be altered significantly by selection of the esterification process. This is particularly true with isophthalic/terephthalic polyesters which are slower reacting than phthalic anhydride. By using a two stage or modified two stage reaction with these aromatic diacids, the molecular structure of the resultant polyester can be changed to markedly improve heat distortion temperature, hydrolytic stability and chemical resistance2.In the two stage process the aromatic acid and glycol are fully or partially reacted before the faster reacting unsaturated acid is added to the cook. This processing method, compared to charging all ingredients at once (one stage method), also leads to a more random distribution of the unsaturation in the polymer chain which changes the character of the final cross linked network in the cured resin. Cure plays one of the most important roles in the chemical resistance developed by unsaturated polyester resins. Theoretically, the curing reaction should go to completion at room temperature with all the double bonds converted to single bonds in the three-dimensional network. However, complete cross linking is rarely achieved at ambient temperatures. This then will result in reduced chemical resistance and, quite often, poorer than expected mechanical properties. In addition, unreacted diluent (styrene ) can remain in the not-so-well cured polymer leading to major problems when the polyester is used for food grade applications. Accordingly, maximum chemical resistance and certain other property improvements can most often be achieved by utilizing elevated temperatures for ‘post cure’ of the polyester resin finished product. Unsaturated polyester resins are used in the manufacture of a broad range of plastic products. A high percentage of these products utilize reinforcing materials, particularly fiberglass. It is estimated that less than 20% of the polyester resins produced are utilized in appli-
cations which do not involve reinforcing materials. These so-called casting applications include buttons, bowling balls, putties, cultured marble, gel coats and decorative products. The marble industry and the more recently developed polymer concrete industry represent outstanding applications for highly filled unsaturated polyesters which offer very economical materials to the building and construction industry. Fiberglass reinforced polyesters (FRP) are used in the manufacturing of boats, automobile and truck parts, building panels, corrosion resistant equipment such as pipes, tanks, ducts, scrubbers, etc., appliances and business equipment, electrical equipment, construction products such as grating and railing, sporting equipment and consumer products that are almost endless. According to the Composites Institute of the Society of Plastics Industry (SPI), automotive, construction, marine and corrosion resistant equipment are the four largest FRP markets, in that order, in the United States which produces 2.5 billion pounds of FRP annually. Mechanical properties are most often the critical factor in the selection of a polyester resin for a specific application. Testing of mechanical properties for both resin castings and fiberglass remforced composites is carried out using standardized ASTM (American Society for Testing and Materials) tests for all plastics. ASTM D-638
Standard Test Method for Tensile Properties of Plastics ASTM D-790 Standard Test Method for Flexural Properties of Plastics ASTM D-695 Standard Test Method for Compressive Properties of Rigid Plastics ASTM D-256 Standard Test Method for Impact Strength (IZOD) of Plastics ASTM D-648 Standard Test Method for Heat Distortion Temperature of Plastics ASTM D-2583 Standard Test Method for Barcol Hardness of Plastics
Polyester resins 39 As mentioned earlier, glycol selection has a produce a rigid polyester which tends to be significant effect on the properties of poly- hard, brittle and lower in tensile elongation. esters. Ether glycols are of great value in Higher unsaturation also leads to higher heat increasing tensile elongation and impact distortion temperature resins. The latter is also strength which is of great importance in auto- achieved by formulating higher molecular motive, casting and liner applications. A weight resins with the chlorendic, bisphenol A principal deficiency of polyester resins is lack and dicyclopentadiene building blocks. As of alkali resistance because the ester linkages expected, all of these resin classes are more are subject to hydrolysis in the presence of brittle and have low tensile elongation. The caustics. Accordingly, increasing the size of the major exception in this scenario are the glycol has the same effect as reducing the con- iso/terepolyesters. Using the multi-stage procentration of attackable ester linkages. Thus, a cessing methods described earlier, these resins resin containing neopentyl glycol, propxylated can be formulated with reasonably high molebisphenol A, or trimethyl pentanediol will cular weights (more linearity) to give very exhibit improved water and chemical resis- tough resins having a good balance of tentance which is highly important in gel coats, sile/flexural properties plus higher tensile corrosion resistant equipment, construction elongation and heat distortion temperatures. Obviously then, when the end use criteria products and many consumer products. The major effect on polyester physical prop- requires the 'something more' than is offered erties is, however, provided by the by general purpose polyesters (orthophthalics unsaturation content in the polyester polymer. and dicyclopentadienes),the formulator turns Higher unsaturation makes for more cross to iso/terepolyesters which have no disadvanlinking and a stiffer cured composite. tages compared to general purpose resins Accordingly, the formulators' selection of other than slightly higher cost. Table 2.1 summarizes the property and unsaturated acid to saturated acid ratio which determines cross linking density can move the application status for the various classes of resin flexural modulus from rigid to resilient unsaturated polyesters. to very flexible. In most cases, a 1/ 1 ratio will Table 2.1 Properties and applications of unsaturated polyesters
Class
Characteristics
Orthophthalics, dicyclopentadiene
Rigid, resistant to crazing, light in color
Isophthalics/terephthalics
Tough, good impact and overall mechanical properties, resistant to environmental elements and moderate chemical attack. Highly resistant to aromatics Rigid, high heat distortion, highly resistant to oxidizing chemical environments Rigid, high heat distortion, highly resistant to most chemical environments particularly caustics
Chlorendic Bisphenol A fumarates
Uses Boats, tub/shower, spas, marble, consumer products, buttons, corrugated sheet, building panels, seating, decorative products Automotive parts, gel coats, electrical, bowling balls, trays, gasoline, tanks, septic tanks, swimming pools, tooling, aerospace products, corrosion, construction products Corrosion resistant tanks, ducting, stacks, industrial vessels Corrosion resistant tanks, piping, stacks, industrial vessels
40 Polyester and uiny ester resins 2.3 VINYL ESTER RESINS
Although vinyl esters have often been classified as polyesters, they should be designated separately because they are typically diesters with a recurring ether linkage provided by the epoxy resin backbone.
Vinyl ester resins are the most recent addition to the family of thermosetting polymers. Although several types of these resins were synthesized in small quantities during the late 1950s, it was not until the mid-1960s that commercialization, principally by Shell and Dow 2.3.1 VINYL ESTER RESIN TYPES Chemical led the push to establish an extremely important segment of today’s com- Aside from the fire retardant versions of vinyl posite industry. Vinyl esters are unsaturated ester resins which are discussed in the next resins made from the reaction of unsaturated section, there are two basic types of vinyl carboxylic acids (principally methacrylic acid) esters having commercial significance. These with an epoxy such as a bisphenol A epoxy are the general purpose lower molecular resin. The typical structure of a vinyl ester weight vinyl esters and the higher heat resistant vinyl esters with greater cross link resin is shown in Fig. 2.8. density. The structure of vinyl ester resins shows several important features which account for the resultant exceptional properties of vinyl General purpose vinyl esters ester resins. There is an epoxy resin backbone with a high molecular weight that provides These are principally methacrylated epoxies excellent mechanical properties combined made by the reaction of methacrylic acid with with toughness and resilience. Secondly, vinyl a bisphenol A epoxy resin. When dissolved in esters display terminal unsaturation which styrene monomer they provide a thermosetting makes them very reactive. They can be dis- resin with good heat resistance, excellent solved in styrene and cured like a mechanical properties (particularly high tenconventional unsaturated polyester to give sile elongation) and outstanding chemical rapid green strength. Obviously, the vinyl resistance to acids, bases, hypochlorites and ester structure also enables convenient many solvents. homopolymerization which could lead to high heat distortion products. Finally, vinyl esters Heat resistant vinyl esters have much fewer ester linkages per molecular weight which combined with the acid resistant These vinyl esters have higher density cross epoxy backbone, give outstanding chemical linking sites available which leads to a more resistance (acids, caustics and solvents) to this heat resistant polymer network. They are proclass of resins. duced from novolac modified epoxy resins
OH H - CI CH2- 0 G
0 7H2 I
c=o
O
CH3
OH O - CH2-CI - H
7% 0 I
c = o
I C-CH3
C-CH3
CH2
CH2
II
Fig. 2.8
T
I
II
Vinyl ester resins 41 and methacrylic acid which provides more cal properties can be 'tailored' to meet the unsaturation sites and higher molecular requirements of specific applications. Another weight due to the epoxy backbone. These unique property of vinyl ester is the bondabilvinyl esters increase the heat resistance by ity of these resins to other surfaces. They are 17-27°C (30-50°F) over the general purpose not as good as epoxy resins in this charactertypes. This often translates to higher useful istic, but obviously the epoxy resin operating temperatures for vinyl ester based component gives them a boost over other reinforced plastics even in corrosive environ- unsaturated polyesters in this area. A case can ments. The higher-density cross linked vinyl also be made for vinyl esters providing better esters are less resilient (lower tensile elonga- fiberglass wet out in FRP composites due to tion) but still have excellent mechanical the backbone hydroxyl groups and their interproperties. Cure of the higher cross linked action with these groups on the fiber surface. vinyl esters may require the use of different Some fabricators have reported that observperoxide catalysts to reduce the peak able resin savings can be achieved with vinyl exotherm and thereby prevent cracking/craz- esters because of this characteristic. However, vinyl esters such as bisphenol A ing in resin rich areas. In other words, resins of this type are more reactive and more caution is polyesters and chlorendic polyesters are made required in the fabrication of FRP laminates. from higher cost materials and often require extended process times which leads to higher finished cost. Accordingly, the specifier/fabri2.3.2 PROPERTIES/APPLICATIONS cator turns to commercial applications where The development of vinyl esters has led to the the improved performance of vinyl esters can fastest growing segment of the thermosetting justify the premium price of the finished comresin industry today. This is not surprising, posite. The foremost application for vinyl esters is since vinyl esters combine inherent toughness with outstanding heat and chemical resis- in glass reinforced laminates for corrosion tance. In all other thermosetting resin types resistant equipment. Because of outstanding one has to sacrifice some heat resistance and chemical resistance combined with excellent often chemical resistance to increase resiliency mechanical properties, vinyl ester based FRP and toughness. Unlike polyesters, vinyl ester tanks, piping, scrubbers, fans and ductwork resins possess low ester content and low are being specified for waste water treatment unsaturation which results in greater resis- plants, mining facilities, chemical processing tance to hydrolysis, lower peak exotherms and storage units, semi-conductor chip operaduring cure and less shrinkage during cure. tions, pulp and paper manufacturing and odor They are easily dissolved in reactive control facilities. Since FRP corrosion resistant monomers such as styrene which provides equipment is the fastest growing segment of easy handling and transportation to the fabri- the US composites industry, the future for cation site. As with polyesters, other reactive vinyl esters looks extremely strong. They are monomers such as vinyl toluene, chlorostyrene comparable to other premium resins for chemand f-butyl styrene can be employed with few ical resistance and secondary bonding combined with a good balance of chemical problems. The toughness of vinyl esters comes from resistance (acids, bases, solvents) at the same the epoxy resin backbone. Since the molecular or lower cost. As a result, chlorendics and weight and structure of the epoxy resin can be bisphenol A polyesters have been reduced to varied like the polyester resin building blocks, 'niche' applications where their specific propphysical properties such as tensile elongation, erty advantages such as heat resistance and heat distortion temperature and key mechani- resistance to oxidizing environments demand
42 Polyester and viny ester resins their use. Since iso/terepolyesters also give an excellent balance of properties in corrosion applications, these unsaturated polyesters and vinyl esters now dominate the corrosion market. The bonus provided by vinyl esters is of course higher heat resistance and extended life at higher operating temperatures, but at significant additional cost compared with the iso/ terepolyesters. The next major market area for vinyl esters utilizes the high tensile elongation characteristics of these resins to produce linings and coating with outstanding adhesion to other types of plastics and conventional materials such as steel and concrete. For example, vinyl esters are an excellent barrier coat for fiberglass boats and acrylic spas. Vinyl ester corrosion coatings are used everywhere today for steel tank linings and industrial flooring. In dual laminate structures, a vinyl ester is often the back up for exotic thermoplastics or the superior corrosion barrier for lower cost polyesters in many FRP tank and pipe applications. The growth of vinyl esters has also been boosted by their excellent handling characteristics and ease of cure. For example, vinyl esters are much preferred by FRP fabricators in filament winding operations because of excellent glass wet out and in fabrication of large structures because the resins are forgiving and provide predictable curing over a wide range of temperatures. The latter situation has resulted in a virtual exclusive use for vinyl esters in field fabrication of large FRP structures.
Table 2.2 summarizes the resin casting properties of the various resins used in corrosion resistant applications today. The outstanding balance of properties provided by vinyl ester resins is obvious and bodes well for continued strong growth in US corrosion markets. Other significant markets for vinyl esters includes pultruded construction and electrical components, automotive structural applications, polymer concrete vessels for mining and chemical operations, grating, high performance marine applications and sporting goods. 2.4 FLAME RETARDANT VERSIONS
The need for flame retardant polymers is essential in many plastics applications today. The combustibility of plastics has drawn so much attention to the safety aspects of these materials in construction applications, that designers and specifiers have been pressured by fire officials to provide fiberglass-reinforced construction materials that exhibit low flame/low smoke characteristics. Since all plastics are based on organic constituents, they are inherently flammable and once ignited will burn until they are completely consumed. There are, however several methods available for making thermosetting resin flame retardant and these provide the capability to supply fire retardant FRP and corrosion resistant/fire retardant FRP for the numerous applications that have a need for some degree of fire retardancy.
Table 2.2 Resins for corrosion resistant applications
General purpose vinyl ester Heat resistant vinyl ester Chlorendic polyesters Bisphenol A polyester Rigid isopolyester Resilient isopolyester
Tensile, psi
Flexural, psi
12 500 13 000 5 500 10 000 8 500 12 500
20 500 20 000 10 000 16 500 19 000 20 000
Elonga f ion break, % 6.7 5.6 1.4 3.2 1.9 4.4
HDT, OF
221 248 284 288 234 201
Flame retardant versions 43 range of chemical environments, both acid and alkali, at operation temperatures similar to the general purpose vinyl esters. Brominated high molecular weight isopolyesters offer economic advantages and are suitable for moderate corrosion applications. These two resin types have become the workhorses for the waste water/odor control FRP market and the chemical and pulp /paper industries because they exhibit excellent impact properties combined with good overall corrosion resistance. 2.4.1 CHEMISTRY AND APPLICATIONS Variations of these resins are used to meet Flame retardancy of unsaturated polyester MIL-R-21607 or MIL-R-7575 requirements. Dibromoneopentyl glycol formulated with and vinyl ester resin is an extension of the nonflame retardant systems (as discussed above). carefully selected chemical building blocks Almost all of these resins can be reformulated provides resins for exposure to severe weatherto include a halogen in the chemical composi- ing conditions. The construction industry uses tion by either blending or by an in situ cook of these resin systems, which are specially formuthe resin. There is an advantage to locking in lated to meet optimum fire retardance for the the halogen in the original resin cook, in order continuous line products of corrugated and flat to chemically tie in the halogen (Cl, or Br) to sheet panels. Such systems are formulated with prevent migration of the halogen when sub- ultraviolet (W) stabilizers and acrylates to jected to thermal degradation. While flame achieve excellent color stability with acceptable retardancy can be achieved with additives low smoke and flame spread (FS)properties. In (Dekabrom or Dechlorine), these additives most cases, these formulations offer good have not been used for high performance chemical resistance for splash and spill on the applications in either the corrosion or con- exposed surfaces. Highly filled halogenated resin systems are designed to accept high filler struction industries (corrugated FRP panels). The chlorendic resins were developed in the loading with aluminum trihydrate (ATH) and 1950s and were based on HET acid (hexa- other synergists to meet DOT requirements for chloroxyclopentadiene). Other formulations low smoke, low flame spread properties followed, based on either tetrabromo bisphe- (ASTM-E-662 and E-162 respectively). Values no1 A (TBBPA) or dibromo neopentyl glycol of <150 smoke and 10 flame spread are (DBNPG). These components react with achieved. Highly filled resins are specified for maleic anhydride or fumaric acid in the pres- applicationswhere people could be exposed to ence of a difunctional glycol, to produce flame indoor fires, such as underground transportaretardant unsaturated polyesters that can be tion. Low smoke allows visual capability to combined with styrene monomer, or other exit an entrapped area. Electrical applications often require the monomers used for smoke control, such as methyl methacrylate. The use of bromine as addition of halogenated base resins to achieve the halogen in the resin building block has flame snuffing properties resulting from high proved to be the most efficient way of achiev- voltage shorting or sparking. Such fire retaring optimum flame retardant thermosetting dant systems are used for compression resins. Certainly a lower percentage of molding of complex electrical shapes, using bromine than chlorine is required for satisfac- BMC or SMC molding materials. Wet mat molding is also used to produce flat sheet for tory reduction of flame spread. Brominated vinyl esters handle a wide electrical insulation components. Flame retardancy can be achieved by using numerous additives, both organic and inorganic. However, most of these have a negative effect on mechanical and/or chemical resistance properties. Accordingly, the most widely used system for achieving optimum fire retardancy will be covered here, namely, halogenated thermosetting resin systems combined with inorganic synergists.
44 Polyester and viny ester resins When compared to the non-flame retardant versions, the addition of a halogen to the resin formulation has very little, if any, effect on the chemical or mechanical properties of a FRP laminate, whereas the use of additives such as aluminum hydrates, clays, carbonates or fumed silica will have a direct and adverse effect on the chemical resistance of FRP laminates. In critical corrosion applications such filler additions could result in early failure of FRP laminates. 2.4.2 TESTING AND CLASSIFICATION
Most halogenated resin systems require a synergist such as antimony usually 3-5'/0 Sb,O, by weight of resin in order to achieve a class I flame spread rating. Other proprietary synergists can be substituted, especially in translucent laminates as used for siding and roofing materials. Antimony oxide enhances the flame spread rating by forming a char on the burning surface of a laminate and effectively subdues the rate of flame progression by snuffing the flame out when the source of ignition is removed or extinguished. Such flame resistant FRP laminates will burn when subjected to a high temperature flame source, but the rate of burning is substantially less than for non-flame retardant systems. The use of Sb,O, will turn laminate opaque, which restricts visual inspections of a laminate in production. It is not a good idea to allow the addition of Sb,O, in the corrosion resistant barrier. Alone, Sb,O, will not improve a non-halogenated resin; its use becomes an unnecessary expensive filler with no flame retardant properties. The most commonly used test method for evaluating flammability is ASTM Method E-84 (the tunnel test) also known as NFPA 255. This test method measures the comparative burning characteristics of a material by evaluating the flame front propagation over the surface of the test material, which is exposed to controlled temperatures in a forced air chamber or tunnel. A flame spread (FS) classification (FSC) is obtained which measures the ignition time
and distance of the flame front advancing down the test tunnel during a ten-minute duration test and compared to those values established for asbestos-cement board (at 0-FS) and red oak material (at 100 FS). This test method establishes the rating at 0-25FS as Class I, 25-75FS as Class I1 and 76-plus FS as Class 111. During the test procedure, the smoke emission is measured and can range from 450 to 1000 or more for unfilled laminates. When additives such as ATH are used, smoke emissions of less than 450 can be achieved. Corrosion resistant FRP ducting exhibiting low flame/low smoke characteristics is required for waste water/odor control and semiconductor applications. This can be achieved with brominated vinyl esters or brominated isopolyesters as the base for the FRP duct which is then coated with an intumescent paint to reduce smoke emissions. Such systems are currently qualified by the International Conference of Building Officials (ICBO) with tunnel test ratings of <25FS and <50 smoke development. Unfortunately, many specifying engineering companies will request and specify values that are not readily achievable for most flame retardant resin systems. Ideally, a Class I system with 25FS (max) to 450 (max) smoke is acceptable for corrosion service, when the use of additives cannot be tolerated. A FSC of <25FS is usually acceptable for most applications. However, some specifiers will claim that 15-20FS is better than 20-25FS, when under actual burning conditions there is a negligible difference in the combustibility of FRP laminates. When selecting a flame retardant resin, it is important to qualify a system to meet the properties required. Flame retardant resins are available in a wide variety of formulations, including lower cost general purpose to premium grade types with better high temperature properties. Class I flame spread thermosetting resin systems can also be achieved without the need for a synergist like antimony oxide. These
Design considerations 45 systems have a higher halogen content which obviously increases resin cost. However, these materials allow for the production of translucent FRP products that are desirable in many construction applications. Such a resin system has been specified by Disney engineers for their architectural applications at Disney entertainment centers. 2.5 DESIGN CONSIDERATIONS
Since most thermosetting resin systems are used with fiberglass reinforcements, it is important to consider material selection and the fabrication process in establishing the design of the FRP composite.
product would essentially ’unzipper’ and fail. The polyester and vinyl ester resins described in this chapter offer a wide selection of materials which will accomplish the need to protect the fiberglass and, at the same time, provide optimum performance properties dictated by the end use application. The first agenda in proper resin selection involves an analysis of the key performance requirements of the end use application. This should be very thorough as follows: 0 0 0 0 0 0
0
2.5.1 MATERIAL SELECTION
0
0
strength requirements; thermal requirements; chemical exposures; electrical requirements; color requirements; surface requirements; environmental exposures; fire resistance needed; smoke requirements; potential upset conditions; number of parts required; life expectancy.
The type of fiberglass reinforcement, place0 ment in the composite and fiberglass content 0 determines the strength of an FRP composite 0 and provides the mechanical properties dictated by the end use requirements. In any part Resin selection is obviously very important in made of FRP, the strength of the part will any FRP application as the above list of design increase directly in relation to the percentage criteria illustrates, but is absolutely vital in of fiberglass in the total weight of the compos- corrosion applications. Corrosive attack on a ite. In addition, the arrangement and type of FRP laminate along with fire is the most critifiberglass will have important effects on the cal situation the composite will face. A resultant physical properties since the strength fiberglass building panel properly made can obtainable in the finished part will be in the perform for an indefinite number of years, but even a properly fabricated FRP tank exposed direction of the fibers. The selection of the thermosetting resin sys- to concentrated acids at elevated temperatures tem will determine the chemical, electrical and may only be good for 10-15 years. The best thermal performance of the FRP product. example of this is in the pulp and paper indusHowever, the most significant contribution by try where vigorous chemical attack on FRP the resin relates to ‘life’ of the composite, equipment can dictate replacement on a rousince the resin must protect the fiberglass. tine basis, say every five years. In spite of this, Accordingly, if the resin fractures or blisters in FRP equipment may still offer the most cost any manner that permits an attack on the glass effective material of construction. Although all FRP composites will be fibers, the composite will lose strength rapidly or delaminate. An interesting way to visualize attacked in the same manner in a particular this is to consider a FRP pipe made by the fila- environment, certain types (chlorendic, ment winding process. If the continuous fiber bisphenol A, vinyl esters and isopolyesters) strands providing the hoop strength to the are significantly more resistant. These then product are severed by chemical attack, the make up the list of corrosion and heat resistant
46 Polyester and viny ester resins resins to choose from in addressing a specific application. A good rule to follow here is that no single FRP resin can handle every kind of environmental problem, so resin selection is of the utmost important. It should also be understood that knowledge of the molecular structure of these higher performance resins does not eliminate the need for actual testing to determine resin suitability in a given application. For example, certain vinyl esters are reasonably resistant to alkaline exposure while other types are poor. The corrosion fabricators’ guide for the suitability of a thermosetting resin in a corrosive environment is ASTM C581. The procedure involves complete exposure of a FRP test laminate for one year, with intermittent strength testing, to establish a curve which depicts loss of flexural strength versus time. It is absolutely essential that the resin selected for that environment form a plateau during the one year test period. Obviously, it is also important that this plateau be achieved at a satisfactory retained flexural strength. Table 2.3 summarizes a comparison of FRP properties of various thermosetting resin types versus carbon and stainless steel. 2.5.2 EFFECT OF PROCESS AND END USE REQUIREMENT
There is an old saying in the FRP business that heightens awareness of the fabrication
process. Simply, ’you can select the best resin and fiberglass in the world and if you don’t put them together correctly - failure will probably result’. Material selection added to design and production requirements leads to a determination of the fabrication process. Many methods of fabrication are used to manufacture products for the numerous FRP markets. These methods vary from hand lay-up/sprayup, filament winding and resin transfer molding which utilize low temperature curing to various high temperature molding compound (SMC), pultrusion, and continuous panel. The designer must analyze the end use property requirements such as color, surface characteristics, strength and chemical resistance requirements and then add-in cost factors, part volume, part size and finishing to finalize the selection of process. For example, transportation body panels would be a high volume application requiring outstanding surface finish and excellent strength properties. All of these can be satisfied with a isopolyester sheet molding compound that is compression molded under heat and pressure. This process can be automated and delivers the highest volume and highest part uniformity of any thermoset molding method. Lower part finishing cost is achievable because subsequent trimming machining is minimized. Corrosion resistant equipment would be fabricated, on the other hand, by either filament
Table 2.3 Comparison of properties of various types of FRP lsophthalic Corrosion resistance Acids Alkalis Peroxides Hypochlorites Solvents Flame retardance Structural strength Thermal insulation A = High, B = Moderate, C = Low
B B C C B C A
A
Orthophthalic Chlorendic Bisphenol A fumarate C C C C C C B A
A C A A B
A A A
A A B B B C A A
Vinyl ester B A B A
B A A A
Carbon steel C B C C A
A A C
Stainless steel B B C C A A A C
References 47 winding or hand lay-up/spray-up processing. The former gives the highest strength to weight ratio of any FRP manufacturing process. However, the most important consideration is that these two process methods allow the easy creation of an effective resinrich corrosion barrier which is mandatory to satisfactory FRP life expectancy in corrosion applications. The purpose of this barrier is to isolate the fiberglass reinforcements from attack that would result in wicking, blistering and delamination. The satisfactory corrosion barrier should be about 125 mils (3175 pm) thick, fabricated with glass or polyester veil on the surface, backed up by 2-3 plies of type E chopped strand mat. The resin-rich corrosion barrier should be constructed with the very best resin available in terms of chemical resistance to the expected environment. For example, the corrosion barrier for 26% hydrochloric acid should employ a vinyl ester and two layers of polyester veil. It is always wise in dual laminate construction (different liner and wall resins) to utilize the material with the higher resiliency (higher tensile elongation) in the liner portion of the laminate structure. A final example of process effect on laminate properties should address the rapidly growing world of FRP pultrusion. This fabrication method provides very high strength due to high fiber concentration and orientation
parallel to the length of stock. Pultrusion is an automated, low labor system which can use any type of thermosetting resin. However, resilient resins such as isopolyesters and vinyl esters are much preferred because of the very high glass content in the finished part. Low cost reinforcement is adaptable to putrusion because the glass weight percentage is high. Pultrusion is used for FRP structural and electrical applications primarily, but the weight and density of the finished product does provide moderate corrosion resistance properties. The value received from good design based on proper selection of materials and process can be very rewarding for FRP composites. The systematic analysis of end use requirements, economic requirements and competitive materials will enable the composite designer/specifier to optimize the cost/performance of FRP as a material of construction. The matrix materials available today give the fabricator sufficient opportunities to meet his final objective of providing the right product at the lowest cost. REFERENCES 1. Ellis, C., US Patent 2 255 313; appl. August 6, 1937. 2. Amoco Chemical Company, Bulletin IP-70a, Chicago, Illinois.
EPOXY RESINS
3
L.S. Penn and €3. Wang
3.1 INTRODUCTION
Epoxy resin systems have achieved acceptance as adhesives, potting compounds, molding compounds and as matrices for continuous filament composites used in structural applications. In this chapter, we discuss epoxy resins in their role as matrices in fiber composites. In this role, they possess several advantages over other types of polymers. The main advantages are: 0
0
0
0
inherently polar nature that confers excellent adhesion to a wide variety of fibers; relatively low cure shrinkage that makes dimensional accuracy of fabricated structures easier to obtain; no volatile by-products of the curing reaction to cause undesired bubble or void formation; crossIinked structure that confers excellent resistance to hostile environments, both aqueous and nonaqueous.
In addition to these advantages, epoxy resins have tremendous versatility because they can be formulated to meet a broad range of specific processing and performance requirements. To know how to take advantage of the formulation options, the engineer needs to have an elementary understanding of epoxy resin chemistry and structure-property relationships. This chapter attempts to provide that by presenting information about the resin system constituents, how they react together to form a Handbook of Composites.Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
crossllnked network and how they lead to different processing parameters and final properties in the formulated system. The chapter also describes the role of cure monitoring and property evaluation in epoxy resin system technology. 3.2 GENERAL DESCRIPTION OF THERMOSETTING SYSTEMS
3.2.1 DEFINITIONS
According to common chemical practice, molecules can be classified by the functional groups they contain. Thus, a molecule containing the epoxide group (shown below) as part of its structure is called an epoxide, regardless of the remaining details of the molecule.
/O\
-c-cI
I
In practice, other types of molecules are added to the epoxide to formulate a thermosetting system, i.e. one that will undergo a curing reaction to harden into a rigid form. The confusing practice has been followed of referring to both the epoxide alone and the formulated system as 'epoxy resin'. One must determine which is truly intended by the context in which it is used. For the purposes of clarity in this chapter, we will use epoxide when referring to the epoxide constituent alone and will
General descripfion of thermoseffingsystems 49 use the term epoxy resin when referring to the uncured or cured formulated system. 3.2.2 THE THERMOSETTING (CURING) REACTION
The thermosetting reaction is the joining of many small molecules by chemical reaction to produce an extended network structure. Although this process is a polymerization, it is distinct from the type of polymerization that forms many individual long chains; the thermosetting reaction unifies all the constituent monomers into a single large molecule extending to the boundaries of the material. Epoxide molecules in the pure state at room temperature normally do not react with each other and can sit for years in a dry container without mutual reaction. The types of chemicals added to the epoxide to effect network formation fall into two categories: curing agents and catalysts. Curing agents, sometimes called hardeners, are added in significant amounts to the epoxide and react with it to become a part of the crosslinked network. These curing agents can be aliphatic amines, aromatic amines, or anhydrides. Catalysts, on the other hand, are added in extremely small amounts to cause the epoxide molecules to react directly with each other, i.e. to homopolymerize. Sometimes the chemicals used as catalysts for homopolymerization can be used for another purpose; when added in small amounts to epoxide-curing agent mixtures, they will accelerate the curing reaction. In this role they are called accelerators rather than catalysts. In the sections below, we discuss the chemical reactions involved in network formation, both when the different curing agents are used and when homopolymerization occurs.
Amine curing In amine curing agents, each hydrogen on an amine nitrogen is reactive and can open one epoxide ring to form a covalent bond 1-5.
When the amine nitrogen contains two hydrogens, each reacts with a different epoxide ring. This scheme is shown in Fig. 3.1, where the developing network is evident. This scheme applies to both aliphatic and aromatic amine curing agents. The reaction between epoxide and amine produces a C-N bond, whose environmental resistance is good, but whose stability to elevated temperature is highly dependent on the adjacent molecular structure.
-
OH I
OH
I
CH-CH,
-
OH I CH-CH2
\
/
cH2-m-
/N-R-y
OH I
CH,-CH-
Fig. 3.1 Reaction scheme for the reaction between epoxide and amine curing agent. Each hydrogen reacts with an individual expoxide group. Thus the primary amine group acts bifunctionally and the secondary amine group acts monofunctionally. The epoxide group acts monofunctionally.
From the reaction scheme, it is obvious that the correct relative amounts of epoxide and amine curing agent must be used. If there is ,an imbalance, unreacted functional groups will be present and the full properties of a complete network will not be developed. The correct amounts by weight to combine are determined by computing the weight of curing agent that contains one chemical equivalent of amine hydrogens and matching that with the weight of epoxide that contains one chemical equivalent of epoxide groups.
50 Epoxy resins
Anhydride curing
functions as a catalyst to speed up the epoxide-curing agent reaction. (Acceleratorswill be In anhydride curing agents, the anhydride discussed later.) Figure 3.2 shows an idealized groups themselves must be cleaved asymmetscheme, where the cleaved anhydride reacts rically to start the r e a ~ t i o n ' ~Initial , ~ . cleavage with an epoxide ring carbon, opening the ring is accomplished with the help of a small in the process. The negatively charged oxygen amount of accelerator,which is a chemical that formed by the opening of the epoxide ring can proceed to react with a different anhydride group, perpetuating the reaction. In this idealized developing network, each anhydride 0 0II I group is bifunctional, i.e. it links to two different epoxide molecules. In practice, the high NR3 (Accelerator) temperatures required for anhydride cure, plus C' the presence of accelerator, provides conditions /I + for some extent of epoxide homopoly0 merization (described later) to take place, making the actual curing reaction much more 00complex than depicted in Fig. 3.2. The reaction I I between epoxide and anhydride produces pri{c=o fC=O marily ester linkages, which have good L += o stability to elevated temperatures and to most L=o hostile environments except bases. + /O\ Not surprisingly, the correct amount of I anhydride curing agent relative to the epoxide must be used to obtain a well developed netCC'O Cc=O work and the associated good properties. The 'C= 0 L += o correct amounts by weight to combine are esti+ mated by examining the reaction scheme and computing the weight of curing agent needed 00I I to react completely with a given weight of c= 0 epoxide. The simplified reaction scheme of Fig. 3.2 proposes that one anhydride group LI o reacts with two epoxide groups and one epoxI 0 0 ide group reacts with two anhydride groups, I I making the number of anhydride groups cono-cH~-c------c-cH*-o I I equal to the number of epoxide groups c= 0 sumed c= 0 consumed in the reaction.
(c=o +
L
+
O
Fig. 3.2 Simplified reaction scheme for the reaction between epoxide and anhydride curing agent. After cleavage, each anhydride group reacts with two epoxide groups and each epoxide group reacts with two anhydride groups. Thus both the epoxide group and the anhydride group act bifunctionally.
Catalytic curing (homopolymerization) The remaining route to formation of a crosslinked network from epoxide molecules requires homopolymerization'4~7.This can be brought about if small amounts of certain Lewis acids or Lewis bases are added. These operate as true catalysts by initiating a selfperpetuating cationic (Lewis acid) or anionic
Constituents used in formulated systems 51 /O\
R3N -I-CH2-m-
+
+R,NCH2-CH-a2I
0/O\
CH2-CH-
+
+ R3NCH2- CH- CH2I
P
cH2-CH-
I
0-
Fig. 3.3 Reaction scheme for the homopolymerization of epoxide. After ring opening, each epoxide group reacts with two other epoxide groups. Thus the epoxide group acts bifunctionally.
(Lewis base) polymerization. A simplified reaction scheme initiated by a Lewis base is shown in Fig. 3.3. Homopolymerization results in the formation of a network of ether linkages, which has outstanding elevated-temperature stability and resistance to hostile environments.
3.3 CONSTITUENTS USED IN FORMULATED
SYSTEMS 3.3.1 EPOXIDES
Although many different types of epoxides are available, only a few are favored for use in matrices in fiber compositess-lO.These are shown in Table 3.1.
Table 3.1 Structures and characteristics of commonly used epoxides
Epoxy equivalent weighf, g/eq
Viscosity at 25"C, Pa s (cP)
Comments
Diglycidyl ether of bisphenol A (DGEBA)
171-177 180-188
3.5-5.5 (3500-5500) 6.5-9.6 (6500-9500)
185-200
10.0-19.0 (10 000-19 000)
450-550
Melting point 65-76°C (149-167°F)
May crystallize on storage. Example: DER 332 (Dow). Contains small amount of higher polymer to prevent crystallization.Examples: DER 330 (Dow) Epon 826 (Shell). Contains small amount of higher polymer to prevent crystallizalion. Example: Epi-Rez 510 (Hi-Tek Polymers), DER 331 (Dow), and GY 6010 (Ciba). n = 2; used in prepregs. Examples: Epon 1001 (Shell)and DER 661 (Dow).
52 Epoxy resins Table 3.1 Continued
Epoxy equivalent weight, g/eq
Viscosity at 25"C, Pa s (cP)
Comments
Diglycidyl ether of bisphenol F (DGEBF)
/O\ CH2- CH- CH2- 0 158-165
5.0-8.0 (5000-8000)
/O\ 0 - CH2 -CH-
CH2
Isomeric mixture that will not crystallize on storage. Example: GY 281 (Ciba).
Polyglycidyl ether of phenol-formaldehyde Novolac
172-179 176-181 191-210
1.l-1.7 (1100-1700) at 52°C (126°F) 20.0-50.0 (20 000-50 000) at 52°C (126°F) 4.0-10.0 (4000-10 000) at 52°C (126°F)
Example: DEN 431 (Dow). Example: DEN 438 (Dow). Example: DEN 439 (Dow).
Polyglycidyl ether of o-cresol-formaldehyde Novolac
200
Melting point 35°C (95°F)
225
Melting point 73°C (163°F) Melting point 80°C (176°F) Melting point 99°C (210°F)
230 235
Used for high-temperature service: R represents chlorohydrins, glycols, and/or polymeric ethers. Molecular weight = 540. Example: ECN 1235 (Ciba). Same as above, but molecular weight = 1080. Example: ECN 1273 (Ciba). Same as above, but molecular weight = 1170. Example: ECN 1280 (Ciba). Same as above, but molecular weight = 1270. Example: ECN 1299 (Ciba).
Constituents used in formulated systems 53 Table 3.1 Continued
Epoxy equivalent weight, g/eq
Viscosity at 25"C, Pa s (cP)
Comments
N, N, N', N', - Tetraglycidyl methylenedianiline
117-133
Used for prepregs. Example: MY 720 (Ciba).
10.0-15.0 (10 000-15 000)
Triglycidyl p-aminophenol /O\
O-CH2-CHI
I /O\
CH2-CH-CH2 95-107
0.55-0.85 (550-850)
/
N\
/O\
CH2-CH-
CH2
CH2
Used extensively for prepregs and adhesives. Example: MY 0510 (Ciba).
Conspicuously, these epoxides all contain aro- achieve cure with amine curing agents fall in a matic rings in their structures. Aromatic rings wide range, from 25°C to nearly 200"C, confer mechanical rigidity and thermal stabil- depending on the chemical structure of the ity to the crosslinked network. It is also worth amine. The first five entries in the table are noting that some of the epoxides in the table aliphatic amines, which can cure epoxides at have two epoxide functional groups, while room temperature or only slightly above. others have three or four or more. Whether a Aliphatic amine-cured systems also tend to network is developed by mixing the epoxide have low glass transition temperatures, T , (the with a curing agent or is developed from epox- temperature at which the mechanical behavior ide alone by catalytic homopolymerization, a changes from rigid to rubbery) and cannot be large number of the molecules in a given for- used in composites that will experience high mulation must be able to react with more than temperature use. two other molecules in order to form a Most other entries in the table are aromatic crosslinked network instead of merely form- amines, whose ring structures confer solidity and mechanical rigidity. These amine curing ing linear chains. agents require elevated temperature cure, but the networks they produce have high glass 3.3.2 AMINE CURING AGENTS transition temperatures and are suitable for Table 3.2 presents commonly used amine cur- use in composites that will be exposed to eleing agents. The temperatures required to vated temperatures in service.
54 Epoxy resins Table 3.2 Structures and characteristics of commonly used amine curing agents
Amine hydrogen equiv. weight, gleq
Viscosity at 25°C (77"F),
Comments
Pa s (cP) Diethylenetriamine (DETA) HzN-CH2-CH2-NH-
20.6
CH2-CH2 -NH2 Available from Dow as DEH 20.
0.00554.0085 5.5-8.5)
Triethylenetriamine(TETA) H2Nf 24.4
(CH2)z-NH
0.020-0.023 (20-23)
CH2-CH2 -NH;! Available from Dow as DEH 24.
Diethylaminepropylamine (DEAPA) CH3- CH2, CH3- CH2 65
50 (5000)
- (CHZ)3- NH2
Available from Union Carbide.
Tetraethylenepentamine(TEPA)
26-27
0.055 (55)
Available from Dow as DEH 26.
Aliphatic polyether triamine (APTA) H2C f-0 - CH2- CH-(CH3)
I CH3 - CH2 - C - CH2 f-0 I
-CH2- CH-(CH,)
H2C f-0 - CH2- CH-(CH3)
77-82
0.072-0.080 (72-80)
NH2 Y
NH2
%NH2
x + y + z = 5.3. Available from Texaco as Jeffamine T 403.
Constituents used in formulated systems 55 Table 3.2 Continued
Amine hydrogen equiv. weight, g/eq
Viscosity at 25°C (77"F), Pa s (cP)
Comments
4,4'-Methylenedianiline (MDA)
50
Melting point 89°C (192")
Available from Ciba as HT 972 and from Pacific Anchor as Ancamine DDM.
rn-Phenylenediamine (MPDA) NH2 I
27
Available from E.I. duPont de Nemours & Company.
Melting point 60°C (140°F)
44' -Diaminodiphenylsulfone (DDS)
0
0
62
Melting point 170-180°C (338-356°F)
Used mainly in prepregs; yields good shelf life and high-temperature properties. Available from Ciba as HT 976.
3,3' -Diaminodiphenylsulf one
0
PIt
.S
I1
0
62
Melting point 174178°C (345-352°F)
Used mainly in prepregs; reacts more slowly than 4,4' analog. Available from Ciba as HT 9720.
56 Epoxy resins Table 3.2 Continued
Amine hydrogen equiv. weight, gleq
Viscosity at 25°C (77"F), Pa s (cP)
Comments
40% MPDA-60%MDA
38
Eutectic mixture. Available from UniRoyal as Tonox 60-40
1.50 (1500)
Dicyandiamide (DICY)
28
Melting point 207-209°C (405408°F)
3.3.3 ANHYDRIDE CURING AGENTS
Table 3.3 presents commonly used anhydride curing agents. Their structures vary widely and some are liquid at room temperature whereas others must be heated to liquefy.
Slow reacting. Used for prepregs. Available from Cytec Ind.
Elevated temperature, typically in the range 100-200°C, is required to achieve cure with anhydride curing agents. The glass transition temperatures of anhydride-cured systems are high.
Table 3.3 Structures and characteristics of commonly used anhydride curing agents
Anhydride equiv. zueigkt, g/eq
Melting point, "C ("F)
Comments
Phthalic anhydride (PA)
0
148
130 (266)
Available from Amoco.
Constituents used in formulated systems 57 Table 3.3 Continued
Anhydride equiv. weight, g/eq
Melting point, "C ( O F )
Comments
Hexahydrophthalic anhydride (HHPA)
154
40 (104)
Available from Pacific Anchor as Ancadride MHHPA and from Ciba as HT 907.
Nadic methyl anhydride (NMA) maleic anhydride adduct of methyl cyclopentadiene
180
Liquid at 25°C (77°F) (0.200 Pa s) 200 cP
Widely used for prepegs. Available from Pacific Anchor as Anhydride METHPA and from Ciba as HY 906.
Dodecenyl succinic anhydride (DDSA)
270
Liquid at 25°C (77°F) (0.200 Pa s) 200 cP
Available from Dixie Chemical and from Humphrey Chemical.
58 Epoxy resins Table 3.3 Continued
Anhydride equiv. weight, g/eq
Melfing point, "C ( O F )
Comments
Chlorendic anhydride (CAI
371
Needs no cure accelerator, but is high melting and hard to handle; good fire retardant. Available from Velsicol.
231-235 (448445)
Trimellitic anhydride (TMA)
,o
\\
0
193
Available from Buffalo Color.
161-1 64 (322-327)
Maleic anhydride (MA) 0
\\
0
100
53 (127)
Available from Amoco.
Succinic anhydride (SA)
/p l0 100
120 (248)
Available from Buffalo Color.
Constituents used in formulated systems 59 Table 3.3 Continued
Anhydride equiv. weight, g/eq
Melting point, "C ( O F )
Comments
Methyltetrahydrophthalicanhydride
166
Liquid at 25°C (77°F) (0.06 Pa s) 60 cP
Available from Lindau.
3,3', 4,4' - Benzophenone-tetracarboxylicdianhydride (BTDA) 0
\\
0
II
0
o/c~cy&o C '
\\
//
0
0
161
221 (430)
Used mainly in powder coatings; when used as minor component in fiber composite matrix, it improves high-temperature properties.Available from Allco.
3.3.4 CATALYSTS FOR
HOMOPOLYMERIZATION Table 3.4 presents some Lewis acids and Lewis bases that have been found effective as catalysts for homopolymerization of epoxides. Certain Lewis acids, such as boron trifluoride (BF,) produce rapid and very exothermic homopolymerization and need to be used in blocked form. BF,, when blocked with monoethyl amine to form the complex BF,MEA, is latent at room temperature and becomes active only above 90"C, the temperature at which the complex separates. The epoxide homopolymerization that occurs above 90°C is rapid, but has controllable release of heat. Epoxides containing blocked Lewis acids as catalysts have been found to be
useful in prepregs that must be stored for some time without cure advancement prior to being used in fabrication of a structural component. 3.3.5 ACCELERATORS FOR CURING AGENT
SYSTEMS Some Lewis acids and Lewis bases can also be used as accelerators in epoxide-curing agent mixtures to speed up a sluggish reaction. They are added in small amounts, only a few weight per cent, empirically determined to give the best results. Most often they are used to speed up the curing reaction in epoxide-anhydride systems. Table 3.4 indicates which catalysts are also used as accelerators and lists additional chemicals used as accelerators only.
60 Epoxy resins Table 3.4 Structures and characteristics of commonly used catalysts and accelerators
Amine hydrogen equiv. weight, g/e9
Melting point, "C ( O F )
Comments
Benzyldimethylarnine (BDMA)
-
Liquid at 25°C (77°F) (0.1 Pa s, 100 cP)
Lewis base used as an accelerator mainly for anhydride mixtures. Avalible from Ciba as DY 062.
2,4,6-Tris(dimethylaminomethyl)phenol
Liquid at 25°C (77°F) (0.3Pa s, 300 cP)
Lewis base used as an accelerator for epoxide anhydride mixtures to provide room-temperature cure. Available from Rohm & Haas as DMP-30 and from Ciba as DY 064.
2-Ethyl-4-methylimidazole (EMI)
CH3- CH2 - C \
Liquid at 25°C (77°F) (4-8 Pa s, 4000-8000 cP)
N H
CH /
Lewis base used as an accelerator for epoxideanhydride mixtures to provide long pot life and good elevated-temperature properties. Available from Air Products as EMI-24.
Constituents used in formulated systems 61 Table 3.4 Continued
Amine hydrogen equiv. weight, g/eq
Melting point, "C ( O F )
Comments
Boron trifluoride-monoethylene amine (BF,MEA) F
I
F -B:NHz-
CH2- CH,
I
F 85-90 (185-194)
Blocked Lewis acid; used as an accelerator for epoxide DDS systems in high temperature service; provides latency. Available from ATOTech.
3.3.6 DILUENTS
For some types of processing the viscosity of the uncured resin system needs to be lowered. When it is not advisable, as in the case of wet filament winding or hand lay-up, to use volatile solvents as diluents, reactive diluents must be used. Table 3.5 lists some acceptable diluents, all low viscosity liquids containing
epoxide groups. Some of these diluents are monofunctional, i.e. have only one epoxide group, so they cannot form crosslinks in the network like bi-, tri- and tetrafunctional molecules do. However, by chemically attaching to the network, the reactive diluents become a permanent and stable part of it. Other diluents are bifunctional and will form crosslinks.
Table 3.5 Structures and characteristics of commonly used, commerciallyavailable epoxide reactive dilutents
Epoxy equivalent weight, g/eq
Viscosify at 25"C, Pa s (cP)
Comments
Butyl glycidyl ether (BGE)
/O\ CH3-(CH2)3 -0 - CH2- CH-CH2
130-149
0.002-0.003 (2-3)
Example: RD-1 (Ciba).
Octyl, decyl glycidyl ether blend
CH3-((CH2)7 215-235
0.005-0.015 (5-15)
-
/O\ 0 - CH2- CH- CH2 Example: Epotuf 37-057 (Reichhold) and Dy 027 (Ciba).
62 Epoxy resins Table 3.5 Continued Epoxy equivalent ___
Comments
Viscosity at 25"C, Pa s (cP)
weight,g/eq ~~
.
_
_
_
_
~
~
p - t - Butyl phenyl glycidyl ether CH3
/O\ 0 - CH2 - CH- CH2
I
-
Example: Epi-Rez 5014 (Hi-Tek Polymers).
0.015-0.30 (15-30)
220-245
Phenyl glycidyl ether (PGE)
/O\ 0 - CH2 -CH-CH2
Example: Heloxy WC-63 (Wilmington Chemical).
0.006 (6)
150
Cresyl glycidyl ether (CGE)
0.005-0.050 (5-50)
170-190
Less volatile than BGE. Example: Epotuf 37453 (Reichhold), and by DY 023 (Ciba).
Diglycidyl ether of 1,4 - butanediol (BDE)
12@140 -~
0.010-0.025 (10-25) _____
~-
Example: RD-2 (Ciba); not a pure compound. ~~~~
~~
~-
General principles of formulation
63
Table 3.5 Continued Epoxy equivalent
weight, g/e9
Viscosity at 25°C’ Pa S ( c ~ )
Comments
~
Diglycidyl ether of neopentyl glycol
CH3 13CL145
Example: AZ epoxy N (AZS Corp.);not a pure
0.005-0.015 (5-15)
compound.
Diglycidyl ether of polypropylene glycol r
1
175-205
0.30-0.060 (30-60)
n = 4. Example DER 736 (Dow).
305335
0.55-0.100 (55-100)
n = 9. Example DER 732 (Dow).
Vinyl cyclohexene dioxide /O\
A CH - C H ~ O ’ W 76
0.20 (20)
3.4 GENERAL PRINCIPLES OF
FORMULATION
Epoxy resin systems must be formulated on a rational basis and the chemical structure of the constituents forms this basis. Many detailed structure-property relationships can be reduced to rules of thumb and some of those we present in this section. The reader should use these with caution, because often a more detailed examination of the chemical structure reveals conflicting trends and experimentation is needed. First, the presence of significant amounts of aliphatic segments in the chemical structure of
Example ERL 4206 (Union Carbide).
the cured network results in lower rigidity and lower Tg as compared to aromatic rings, or even saturated rings. Thus, aromatic aminecured systems or homopolymerized systems both have high stiffness and high Tg. whereas an aliphatic amine-cured system will have a lower stiffness and lower T,. In terms of processing and cure, the flexibility and mobility of aliphatic segments imparts low viscosity (if a liquid) or low melting point (if a solid). Constituents whose structures are mainly aliphatic react rapidly at room temperature. Thus, an aliphatic amine-cured system is recommended over an aromatic amine-cured
64 Epoxy resins system if low viscosity for processing and a room temperature cure are needed1’,12. Constituents with aromatic ring structures react sluggishly or not at all at room temperature13J4.Thus aromatic-amine cured systems require elevated temperature cures. As cited earlier in the discussion on cure reactions, the relative amount of curing agent to epoxide is important in achieving a welldeveloped network. The formulations presented in Table 3.6 are approximately what would be used for some specific epoxide-curing agent formulations. The reader should verify the correctness of the formulations by computing the appropriate weight ratios from the molecular structures given in Tables 3.1, 3.2 and 3.3. Table 3.6 Formulations for selected epoxy resin systems
Constituents
Epoxide Curing agent PbV
Pbw‘
Diglycidyl ether of bisphenol A triethylene tetramine
100
14
Diglycidyl ether of bisphenol A rnetaphenylene diamine
100
16
Diglycidyl ether of bisphenol A hexahydrophthalic anhydride
100
90
Parts by weight
Finally, through the chemistry of the constituents and the network they form, the formulation d u e n c e s the environmental resistance of the resin. Thus the formulation must be selected with the future environmental exposures in mind. Figure 3.4 compares the environmental resistance of four important types of epoxy resin systems: aliphatic amine-cured,aromatic aminecured, homopolymerized epoxide (BF,MEA catalyst) and anhydride-curedI5. The excellent resistance of the crosslmked network is evident
in all but a few cases; the anhydride-cured system is degraded in strong base due to basic hydrolysis of its ester linkages and both the anhydride-cured and the homopolymerized systems are vulnerable to swelling by the strong solvent, trichloroethylene. 3.5 PROCESSING CONSIDERATIONS IN EPOXY RESIN COMPOSITES
The goals of the processing procedures used to make a good quality fiber composite are to ensure that the resin forms a void-free continuous phase, surrounds each filament, is evenly distributed, is present in the desired amount relative to the fiber and is fully cured. The ease with which these goals can be achieved is highly dependent on the rheological properties of the resin as it progresses through its cure and on the engineer’s ability to evaluate rheology and degree of cure throughout the processing cycle. In the early years of epoxy resin technology, processing procedures were developed by a trial and error approach. This is still a viable approach, especially for engineers skilled in the art of thermoset composite processing. However, in recent years, rheologica116-18 and chemorheologica119~20 models that relate viscosity, rigidity and degree of cure to time and temperature have been developed. These models predict rheological changes during cure and can serve as an aid to processing and cure cycle development. 3.5.1 PROCESSING VARIABLES EARLY IN CURE
One of the major concerns in the early stages of composite processing is resin viscosity. Sometimes, the freshly mixed, uncured resin formulation is inherently fluid enough to penetrate the fiber bundle and surround each filament. (Tables 3.1, 3.2 and 3.3 give viscosity values for several epoxides and curing agents.) Often, however, the viscosity of the uncured resin system is too high and must be
Processing considerations in epoxy resin composites 65
100
75
50
ae I
l m n .-C
25
c1 VI
9 $
0 Sodium hydroxide, 50% 82°C (180°F)
0
E+
Sulfuric acid, 25%
82°C (180°F)
Hydrochloric acid, 25% 82°C (180°F)
L
0 C
.-c0 C
W
w
E
-c3
c 0 100
i
-E 2
; 75
W
U
50
25
0 Distilled water 54°C (13CPF)
Trichloroethylene 54°C (130°F)
Sodium hypochlorite, 6% 54°C (130°F)
Exposure conditions
Fig. 3.4 Environmental resistance of common cured epoxy systems as indicated by flexural modulus retention after environmental exposure Is. Aliphatic amine-cured (TETA),
Homopolymerized (BF,MEA),
Aromatic amine-cured (MPDA), @
Anhydride-cured (PA),
reduced to achieve the desired flow requirements. The two major approaches to reducing viscosity are thinning the mixture with low viscosity organic solvents and adding low viscosity reactive diluents, such as those presented in Table 3.5. Thinning with organic solvents, although
the simplest approach, can result in bubble or void formation within the composite if the solvent cannot escape completely. This could be a problem for component fabrication by wet filament winding, where layers containing the freshly mixed epoxy resin system are placed sequentially on top of one another.
mrm
66 Epoxy resins Filament winding processors do not use solvents to reduce viscosity, rather selecting lower viscosity resins, reactive diluents, or diluting with heat. On the other hand, the use of small amounts of solvent to reduce resin viscosity during fabrication of prepreg (pre-impregnated fiber) presents no problems, since prepreg is made in the form of single, thin-layer sheet, tape or tow from which solvent can vaporize easily and the prepreg is heat treated to eliminate solvent and to advance cure.
-
A
'T
1100
-
- -m
IS00
i
.-
ln
0
1100
, 0
.ln
1100
1000
eo0 800 700
600 500
400 $00
zoo 100
-
-
-
-
-
By contrast, reactive diluents, being themselves epoxides, chemically react to become a permanent part of the crosslinked network. Figure 3.5 shows the relation between L'1SCOS' ity and the amount of diluent added to a viscous epoxide. Ideally, the engineer wants to use just enough diluent to lower viscosity as needed without dramatically altering the properties of the final cured network. The length of time that an epoxy resin formulation remains fluid is important. Liquid-like flow becomes impossible once the gelation stage, marked by an abrupt increase in viscosity, is reached. Time to gelation is called gel time, or sometimes pot life. Aliphatic amine curing agents produce pot lives of the order of minutes or a few hours, while aromatic amine curing agents produce pot lives of 24 h or more13,22,23. Anhydride curing agents typically produce very long pot lives (e.g. two months for NMA) when mixed with epoxides. This is because, as already mentioned, the anhydride group is not very reactive with epoxides unless it is cleaved with the aid of an accelerator molecule. Once the accelerator is activated, the pot life of the mixture will be shortened to a few hours. Pot life can be controlled over a wide range by careful use of accelerators. Each resin formulation has a unique chemistry that imparts a set of processing variables with unique values. Standard laboratory test methods for processing variables that are important early in cure are described by the American Society for Testing and Materials, Philadelphia, Pennsylvania.ASTM numbers of the procedures for determining resin viscosity, gel time and melt flow are listed in Table 3.7.
I l l l l l r l l l l Table 3.7 Laboratory test methods for measuring
0
10
20
30
40
SO
60
70
80
90
IOC
DILUENT, 'Io
processing variables
Processing variable Fig. 3.5 Viscosity versus per cent of diluent in epoxide-diluent mixturesz1.The epoxide is DGEBA and the diluent is BDE (diglycidyl ether of 1, 4-butane diol).
-
Viscosity time (pot life) Melt flow
Reference ASTM D2393 ASTM D2471 ASTM D3795, ASTM D4473
Processing considerations in epoxy resin composites 67 3.5.2 MONITORING OF CURE
Once the cure is underway in an epoxy resin system, all of the properties of the system change rapidly until the final crosslinked network is reached. Not only is flow decreasing and rigidity increasing, but all other properties (electrical, chemical, optical, etc.) are changing. When the curing epoxy resin system is subjected to temperature and pressure changes, the resin response will be characteristic of its degree of cure at that moment. Cure cycle events, such as temperature and pressure changes, need to be carefully timed with this in mind. The degree of cure can be defined in terms of any one of a large number of chemical or physical (including mechanical) properties that change continuously during the curing reaction and reach a constant value at end of the cure. Evaluation of the degree of cure is usually based on one of these properties and full cure is then defined as the point at which this selected property reaches a constant value. Originally, the only available methods for evaluating cure were off-line laboratory meth-
ods where samples were taken and tested at intervals in the cure cycle. The time needed to develop one data point depended on the particular laboratory method. Off-line testing has been made more convenient with the development of continuous monitoring techniques (i.e. data points in real time) used on small dedicated specimens of resin or prepreg. Recently, real time, in situ methods where cure can be monitored in the composite structural component itself have been developed.
Traditional off-line methods Intermittent off-line methods include chemical titration of the unreacted epoxide groups presentz4jz5, specific gravity to measure densificationzh and differential scanning calorimetry to measure the residual cure e~otherm*"~ Figure ~ ~ ~ ~3.6 * .shows a plot of specific gravity data obtained on specimens cured for increasing lengths of time at a single temperature. Figure 3.7 shows differential scanning calorimeter scans for two epoxy resin specimens of the same formulation, but with different degrees of cure.
1.24 1.22 1.20 -
El
\
s c .-
+ $$ + s t + + $
+
+
+
+
1.181.16 -
v)
1.14-
Fig. 3.6 Specific gravity compared with cure time for an aromatic amine-cured epoxy system cured at 120°C26. Volume reduction (densification) during early network formation is rapid and levels off as cure nears completion.
Exotherm
A
.
I
I
I
I
I
1
I
Energy
calls B 1
Endotherm
.
I
I
I
I
I
I
I
Fig. 3.7 Differential scanning calorimeter scans for an epoxy resin system28.Scans for two different specimens, each having a different original degree of cure, are shown, with scan A displaced upward from scan B for graphical visibility. The height of the residual cure exotherm is inversely related to the original degree of cure. The higher exotherm peak in B indicates an original degree of cure lower than in A.
Continuous off-line methods include infrared s p e c t r o s ~ o p y ~ ,parallel ~ ~ , ~ ~ , plate-type bulk diele~trornetry~’,~~ and dynamic mechanical s p e ~ t r o m e t r y ~Figure ~ ” ~ . 3.8 shows the results of infrared monitoring for five different neat resin specimens, each cured at a different temperature and Fig. 3.9 shows data obtained by bulk dielectrometry.
The off-line methods, real time or not, are useful for developing a cure schedule for a new epoxy resin formulation, for optimizing processing variables and for quality control of incoming resins or prepregs. They have also been used successfully for the development of mathematical models of cure kinetics. However, because they are off-line, they cannot be used for process control.
Fig. 3.8 Degree of cure compared with time for an aromatic amine-cured epoxy systemz8.Data for five specimens, each cured at a different temperature, are presented.
Property data for cured epoxy resin systems 69 200
I
I
1
f
I
f
.1
1
f
Decreasing molecular dipole mobility
I
I \
0
I
I
I w
E
- t
Temperature profile
150
increasing molecular dipole mobility er; elevation -
100
%r C 0 .-c
-------I-
P
E
m
.-n
F
.-
v)
v)
50
No further cure at temperature
I
0 0
2
4 Time - h
1
6
I
8
Fig. 3.9 Bulk dielectrometry data compared with time for an aliphatic amine-cured epoxy system28. Dissipation factor (dashed line) is inversely related to rigidity in the developing network. The applied temperature is shown by the solid line.
Modern in situ methods
In situ methods for cure monitoring are real time methods that require sensors small enough to implant and leave in the composite itself. When the information from in situ monitoring is used as continuous input to an appropriate process model, it can be used in process control loops that adjust the processing conditions automatically. One example of an in situ method is low frequency dielectrometry using a very small assembly of interdigitated electrodes called a fringe field sensor. This technique measures changes in ability of permanent dipoles within the resin chemical structure to align themselves with the applied oscillating electric field and also measures changes in the mobility of ions present as impurities in the resin3941.Both of these quantities correlate with resin viscosity in the early stages of cure and with mechanical rigidity in the later stages of cure39,40,42.
Other examples of in situ monitoring are based on fluorescence spectroscopy of tag molecules in the resin43,44 and on infrared45,46 and Raman47spectroscopies of the resin molecules themselves. In these methods, optical fibers are the sensors that transmit the appropriate wavelength of light into the curing resin and also transmit the spectral information back The spectral changes relate directly to the chemical changes that occur as the curing reaction progresses to completion. 3.6 PROPERTY DATA FOR CURED EPOXY
RESIN SYSTEMS
Property data for cured, unreinforced epoxy resin systems are needed for two purposes. First, they are useful when selecting the best fiber and matrix combination for a particular application. Resin system choices can be rapidly narrowed down to a few alternatives
70 Epoxy resins
when comparisons of key properties are made from existing data tables or manufacturers’ data sheets. Second, epoxy resin data are required in micromechanics computations of composite properties. Elastic constants, thermal expansion coefficients, moisture absorption coefficients,
and many other properties of the composite can be computed in advance, if one has the corresponding values for the fiber and the matrix. Tables 3.8, 3.9 and 3.10 present property data for three major resin systems: aliphatic amine-cured, aromatic amine-cured and anhydride-cured. Property data for commonly used
Table 3.8 An aliphatic amine-cured epoxy resin system, room-temperature curable1228 Resin system constituents Epoxy: DGEBA, eg., DER 332 (Dow) Curing agent: APTA, e.g. Jeffamine T-403 (Texaco)
Parts by weight 100 45
Cure cycle: 16 h at 60°C (140°F)for improved properties over room temperature cure Viscosity at 25°C (77°F) 0.8 Pa s (800 cP)
Heat distortion temperature at 1820 kPa (264 psi)
Density of cured resin at 25°C (77°F) 1.16 g ~ m - ~
Coefficient of linear thermal expansion from 298 to 374 K 66x
Volumetric shrinkage After gelation After cure
Average specific heat from 286 to 367 K 1.75~10~Jkg-lK~~
4.4% 4.4%
Water absorption, wt. gain after 2 h in boiling water 0.75%
l ’
a
Impact strength (Izod notched bar test)
11.0 J m-’ of notch
Shear properties Failure stress Modulus of elasticity
61 MPa (8.85 ksi) 1.27 GPa (184 ksi)
Tensile properties Modulus of elasticity
3.24 GPa (470 ksi)
I
120
I
I
“C-’
Coefficient of thermal conductivity At 298 K 0.133 W m-’ K-’ At 318 K 0.174 Wm-’ K-’ At 336 K 0.210 W m-l K-’
a a
Compressive properties Modulus of elasticity 3.48 GPa (504 ksi)
I Cures:
r
62°C (144°F)
16 h @ 60°C
16
._
4
Y m
0
0
1
2
3
4
5
Tensile strain %
0
0.8
1.6
3.2
2.4
Compressive strain % ~~
a
Cured for 24 h at 60°C (140°F) + 24 h at 77°C (171°F).
4.0
-~
~~~
__
Property data for cured epoxy resin systems 71 Table 3.9 An aromatic amine-cured epoxy resin systemI3 Resin system constituents Epoxy: DGEBA, e.g. Epon 826 (Shell) Diluent: BDE, e.g. RD-2 (Ciba-Geigy) Curing agent: MDA-MPDA eutectic, e.g. Tonox 6 0 4 0 (UniRoyal)
Parts by weight 100 25 29
Cure cycle: 3 h at 60°C (140°F) + 2 h at 120°C (248°F) Viscosity at 25°C (77°F)
Water absorption, wt. gain after 6 h in boiling water
1.2 P a s (1200 CP)
I
I
- 14 - 10
60
-
40
- 6
._
8
In VI
E
5
1.54 x lo3J kg-' K-' 1.71 x lo3J kg-'K-'
Coefficient of thermal conductivity 0:243 W m-' K-' At 325 K 0.244 W rn-' K-' At 356 K At 389 K 0.256 W m-'K-' Shear properties Failure stress
80
h
Specific heat At 363 K At 424 K
2.68 GPa (389 ksi) I
121°C (250°F)
Coefficient of linear thermal expansion 6.81 x 10-5oc-1 from 298 to 3755 K
3.7% 5.4%
Tensile properties Modulus of elasticity
H
Heat distortion temperature at 1820 kPa (264 psi)
1.15 Mg m-3 1.21 Mg m-3
Volumetric shrinkage After gelation After cure
I
130°C (266°F)
23 h
Density at 25°C (77°F) Uncured Cured
100
Glass transition temperature
6h
Time to reach 2.0 Pa s Gel time for a 30-g mass at 25°C (77°F)
0.93%
% VI
VI
E
Compressive properties Maximum stress Strain at maximum stress Modulus of elasticity
52 MPa (7.54 ksi) 111MPa (16.1 ksi) 8.0 % 2.9 GPa (420 Ksi)
G
- 4
20 - 2 0 0
2
4
6
a
io-
Strain YO
epoxy resin formulations are often available from resin suppliers. Data for new or unusual formulations must be generated by the user. Whether the data are generated by the resin supplier or the user, it is important that standard test procedures be followed. This will ensure that the resin systems can be compared on an equal basis. Where standardized test
procedures exist, e.g. as from the American Society for Testing and Materials, they should be followed. Where they do not exist, literature references are helpful. Table 3.11 lists some commonly tested properties and the standard methods (American Society for Testing and Materials) describing the tests.
72 Epoxy resins Table 3.10 An anhydride-cured epoxy resin systemI5 Resin system constituents Epoxy: DGEBA, e.g. Epon 828 (Shell) Curing agent: NMA Accelerator: BDMA
Parts by weight 100 90 1
Cure cycle: 3 h at 120°C (248°F) + 24 h at 150°C (302°F) 1.78 P a s (1780 cP) Viscosity at 27°C (81°F) 556 days Time to reach 100 Pa s (1000 cP) 4-6 days Pot life of a 500 g mass at 23°C (73°F) 121°C (250°F) Heat distortion temperature at 1820 kPa (264 psi) Solvent absorption, wt. gain After 24 h in boiling water After 3 h in boiling acetone Tensile properties Maximum stress Strain at maximum stress Modulus of elasticity
0.67% 1.9% At 23°C (73°F) At 100°C (212°F) 72.4 MPa (10.5 ksi) 46.2 MPa (6.70 ksi) 2.7% 7.2% 1.38 GPa (200 ksi) 3.45 GPa (500 ksi)
Table 3.11 Standard test methods for cured epoxy resin systems
Property
___
ASTM Standard Method
Physical and chemical properties: Specific gravity Chemical resistance Water absorption Light and water exposure
D792 D543 D570 D1499
Electrical properties: Volume resistivity Surface resistivity Dielectric strength Dielectric breakdown voltage Permittivity, dielectric constant Dielectric loss
D257 D257 D149 D149 D150 D150
Thermal properties: Heat deflection temperature Glass transition temperature Coefficient of linear thermal expansion Coefficient of thermal conductivity
D648 D4065 D696 C177
Mechanical properties: Tensile modulus and strength Compressive modulus and strength Flexural modulus and strength Impact resistance
D638 D695 D790 D256
Dynamic mechanical properties: Storage modulus Loss modulus Transition temperature
D4065 D4065 D4065
References 73 17. Lee, W.I., Loos, A.C. and Springer, - - G.S. ’Heat of reaction, degree of cure and viscosity of 1. Lee, H. and Neville, K. Handbook ofEpoxy Resins, Hercules 3501-6 resin,’ J. Compos. Mater., Classical Re-issue, McGraw-Hill, New York, 1982.16.510-520. , , 1982. 18. Springer, G.S., ‘Compaction and consolidation 2. Goodman, S’H. ‘Epoxy Resins’y ch’ in of thermoset and thermoplastic composites,’ Handbook of Thermoset Plastics, (S.H. Goodman, Compos. Polym,,1990, 3, 155-175, Ed.), Noyes Publications~ Park New 19. Mijovic, J. and Ott, J. ’Modeling of chemorheolJersey, 1986, pp. 132-182. ogy of an amine-epoxy system of the type used 3. Tanaka, Y. and Bauer, R.S. ’Curing Reactions,‘ in advanced composites,’ J. Compos. Mater., ch.3 in Epoxy Resins: Chemistry and Technology, 1989,23,163-185. Dekkerr In‘., New 20, Mijovic, J. and Wang, H,T, ‘Cure kinetics of neat 1988, pp. 285463. and graphite-fiber-reinforced epoxy formula4. Fischerf M.y F. and Schmid7 R’ tions,’ J. Appl. Polym. Sci., 1989,37,2661-2673. ’Struktureller Aufbau and Physikalisches 21. Product data sheet for Araldite RD-2, CibaVerhalten venetzter Epoxiharze,‘ Makromol. Geigy, Ardsley, New York, 1979. Chem., 1980,181,1251-1287. 22. Penn, L.S. and Chiao, T.T. ‘A long pot-life epoxy 5. Dusek* R. and B1ehar M. ‘Curing Of system for filament winding,’ in Proceedings of resins: model reactions of curing with amines,’ J. the 7th National SAMPE Technical Conference, POlym. SCi., Polym. Chem. Ed., 1977~15,2393-2400. ~October ~ 14-16,i Albuquerque, N~~ ~ 6. Fisch, W. and Hofmann, W. ’ m e r den 1975, vel. (1975), pp. 177-187. Hartungsmeehanismu. der AthexY1i*arze,’ J. 23. Chiso, T.T., Jessop, E.S. and Perm, L.S., Polym. Sci. 1954,12,497-502. ‘Screening of epoxy systems for high perfor7. Morgan, R.J. and Mones, E.T. ’Cure reactions, mance filament winding applications,~ in response Of netWork structure,and Proceedings of the 7th National S M P E Technical diamine sulfonecured tetraglycidyl Conference, Albuquerque, New Mexico, October 4,4‘-diamino methane epoxy matri1&16,1975, Val. 7 (1975), pp. 167-175. ces,’ J. Appl. Polym. Sci., 1987,33,999-1020. 24. Jay, R.R. ’Direct titration of epoxy compounds 8. Resins and Hardeners Product Guide, Ciba-Geigy ~(-hem,~ 1964, 1 36,, 667468. and Aziridines,r ~ Corporation, Hawthorne, New York, 1990. 25. Jahn, H. and Goetzky, P.’Analysis of epoxides 9. Epoxy Resin Manual: Industry Edition, Dow and epoxy resins,, ch. 13 in E~~~~ ~ ~ Chemical Company, Midland, Michigan, 1988. Chemistry and Technology, (C.A. May, Ed.), 10. Specification Guide for Epon and EpOnol Epoxy Marcel Dekker, Inc. New Y’rk, 1988, pp. Resins and Epon Curing Agents, Shell Chemical 1049-1087. Company, Houston, Texas, 1990. 26. Chou, C.T. and Penn,L.S. ’Mechanism by which 11. Chiao/ T.T.r JessoP, E.S. and NeweYf H.A. ‘A an orthocarbonate reduces residual stress in a composite,’ J. Compos. Mater., 1992,26,171-184. moderate- temperature-curab1e epoxy for advanced composites,’ SAMPE Quart., 1975, 6, 27. Carpenter, J.F. and Bartels, T.T., 3842. ’Characterization and control of composite 12. Chiao, T.T. and Moore, R.L. ‘A roOm-temperaprepregs and adhesives,’ in Proceedings ofthe 7th ture-curable epoxy for advanced fiber National SAMPE Technical Conference, Albuquerque, New Mexico, October 14-16, composites,’ 29th Ann‘ PlasticslComposites Institute, SPI, Washington, 1975, vel, (1975),pp. 43-52. D.C., February 5-8,1974, section 16-B. 28. Penn, L.S. and Chiao, T.T. ‘Epoxy resins,’ ch. 15 13. Chiao, T.T., Jessop, E.S. and NeweY H.A. ‘An in ~ ~ n d b o oof kComposites, (George Lubin, Ed.), epoxy system for Van Nostrand Reinhold, New York, 1982, pp. Quart., 1974, 6,28-32. 57-88. 14. Product data sheets, Ciba-Geigy, Hawthorne, 29, Damenberg, H. ‘Determination of functional New York, 1993. groups in epoxy resins by near infrared spec15. Epoxy Resins for Casting, Shell Chemical troscopy; SPE pans., 1963,3, 78-88. Company, New York, 1967. 30. Morgan, R. ’Structure-property relations of 16. Springer, G.S., ’Resin flow during the cure of epoxies used as composite matrices,’ Adv. fiber reinforced composites,’ J. Compos. Mater., Polym, sei,, 1985, 72, 143. 1982,16,400410. REFERENCES
‘Onf.I
~
~
~
i
74 Epoxy resins 31. Yalof, S. and Wrasidlo, W. 'Crosschecking between dielectric measurements, DTA, and other methods of thermal analysis in research and production,' J. Appl. Polym. Sci., 1972, 16, 2159-21 73. 32. Delmonte, J. 'Electric properties of epoxy resins during polymerization,' J. Appl, Polym. Sci. 1959, 2, 108-113. 33. Lewis, A. 'Dynamic mechanical behavior during the thermosetting curing process,' SPE Trans., 1963, 3,201-212. 34. Kreahling, R.P. and Kline, D.E. 'Thermal conductivity, specific heat, and dynamic mechanical behavior of diglycidyl ether of bisphenol A cured with m-phenylene diamine,' J. Appl. Polym. Sci., 1969, 13,2411-2425. 35. Arridge, R. and Speake, J. 'Mechanical relaxation studies of the cure of epoxy resins: 1. Measurement of cure,' Polymer, 1972, 13, 443449. 36. Kline, D.E. 'Dynamic mechanical properties of epoxy resins during polymerization,' J. Appl. Polym. Sci., 1960,4, 123. 37. Babayevsky, P. and Gillham, J. 'Epoxy thermosetting systems: dynamic mechanical analysis of the reactions of aromatic diamines with the diglycidyl ether of bispenol A,' J. Appl. Polym. Sci., 1973,17, 2067-2088. 38. Enns, J.B. and Gillham, J.K. 'The time-temperature transformation (TTT) cure diagram: modeling the cure behavior of thermosets,' J. Appl. Polym. Sci., 1983,28,2567-2591. 39. Senturia, S.D. and Sheppard, N.F. 'Dielectric analysis of thermoset cure,' Adv. Polym. Sci., 1986, 80, 1 4 7 .
40. Day, D.R., Lewis, T.J., Lee, H.L. and Senturia, S.D. 'The role of boundary layer capacitance at blocking electrodes in the interpretation of dielectric cure data in adhesives,' J. Adhesion, 1985,18,73-90. 41. Micromet Instruments, Inc., Cambridge, Massachusetts, Technical literature, 1991. 42. Ciriscioli, P.R. and Springer, G.S. 'Dielectric cure monitoring - a critical review,' S A M P E J., 1989, 25, 3542. 43. Sung, C.S.P., Pyun, E. and Sun, H.L. 'Characterization of epoxy cure by UV-visible and fluorescence spectroscopy: azochromic labeling approach,' Macromolecules, 1986, 19, 2922-2932. 44. Sung, C.S.P. and Mathisen, R. 'Cure characterization of an epoxy network by fluorescence behavior of trans-diaminostilbene,' Polymer, 1987,28,941-945. 45. Compton, D.A., Hill, S.L., Wright, N.A. et al. 'In situ FTIR analysis of a composite curing reaction using a Mid-Infrared Transmitting Optical Fiber,' Appl. Spectros. 1988,6,972-979. 46. Young, P.R., Druy, M.A., Stevenson, W.A. and Compton, D.A. 'In situ composite monitoring using infrared-transmitting optical fibers,' S A M P E J., 1989,25,11-16. 47. Myrick, M.L., Angel, S.M., Lyon, R.E. and Vess, T.M. 'Epoxy cure monitoring using fiber-optic Raman spectroscopy,' S A M P E J., 1992, 28, 3742. 48. Mijovic, J., Kenny, J.M., Nicolais, L. and Pejanovic, S. 'Present and future trends in insitu monitoring of processing of advanced composites,' SAMPE J., 1992, 28, 3946.
HIGH TEMPERATURE RESINS
4
Hugh H.Gibbs
4.1 INTRODUCTION
too much from their high temperature mechanical properties and thermal-oxidative stability. The high temperature resins discussed in this As a result of a extensive work on the part of chapter are defined as a family of aromatic polymer scientists, mostly within the USA, a polyimides having glass transition temperawide variety of products has been developed tures ( T J greater than 316°C (600°F). Other possessing various trade-offs between their resin systems such as the bis-maleimides and processibility and properties. In writing this various aromatic thermoplastics (including chapter the author has attempted to explain the lower Tg thermoplastic polyimides) are disevolution of high temperature polyimides as cussed in Chapters 5 and 6 respectively. matrix resins for advanced composites together Over the years it has been found that the key to achieving outstanding high temperature with a description of their processing charactermechanical properties and thermal-oxidative istics (where available),physical properties and stability is to have a polymer made with aro- long term high temperature performance. matic heterocyclic repeat units where there is a minimum aliphatic content (e.g. aliphatic C-H 4.2 CONDENSATION POLYIMIDE and C=C groups). Such groups can contribute CHEMISTRY to thermal-oxidative instability. Although many types of aromatic heterocyclic polymers 4.2.1 GENERAL COMMENTS ON are possible, one type, polyimides, has turned CONDENSATION POLYMERIZATION out to be the most commercially successful. The highly aromatic character achievable in Studies carried out in DuPont in the 1950s and such polymers is the reason behind their ther- 1960s established that polyimides can be premal-oxidative stability. Also, provided that pared by the reaction of an aromatic flexibilizinglinkages in the monomers are kept dianhydride and an aromatic diamine in a to a minimum, the mherent rigidity of the polar solvent such as dimethyl acetamide or N(NMP). This is repeat units results in the high T which is methyl-2-pyrrolidone essential if an adequate level of hi& tempera- illustrated in Fig. 4.1 for the polyimide based ture mechanical property retention is to be on pyromellitic dianhydride (PMDA) and 4,4'-oxydianiline (ODA). The intermediate achieved. Over the past 25 years various strategies polyamide acid solution is the basis for have been developed to introduce processibil- DuPont's Pyre ML@wire enamel. During the ity into aromatic polyimides without detracting second step, the imidization of the polyamide acid, 2 moles of water are eliminated per repeat unit. Two other DuPont products, Kapton@ ilm and Vespel@SP polyimide preHandbook of Composites.Edited by S.T. Peters. Published polyimide f cision parts are also based on this chemistry. in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
76 High temperature resins
ti
ti
i
4.4-Oxydianiline (ODA)
Pyromellitic Dianhydnde (PMDA)
+
N
:
~
~
II
0
; / N
~C ' o
0
0
PMDAiODA Polyamide Acid
n
Heat (-2H,O) (-solvent)
+
0
n
PMDAiODA Polyimide
Fig. 4.1 Typical reaction sequence for a polyimide from a dianhydride and a diamine.
It turns out that the particular polyimide temperatures in excess of 100°C (212°F) are shown in Fig. 4.1. is unsatisfactory for the pur- reached. Thus high solids monomeric solupose of making prepregs for high temperature tions are possible which are ideal for composite parts. For one thing the prepregging. If, on the other hand, the dianhyPMDA/ODA polyimide is too intractable, dride is the commercially available starting having no detectable T , or melting point material then it must be converted in situ to below its decomposition temperature which is the open ring diester diacid form by prereactwell in excess of 500°C (932°F).In addition, the ing with an alcohol such as ethanol or polyamide acid solutions which are immedi- methanol. ately generated by dissolving the diamine in a suitable solvent and then adding the reactive 4.2.2 CHEMISTRY OF SKYBOND' dianhydride, are unsuitable for prepregging. Ideally monomeric solutions are preferred that Skybond from Monsanto is a product which possess modest viscosities even when the has been available commercially since the solids contents are in the 50-60 wt.% range. mid-1960s. The chemistry of Skybond is illusThis desired combination of properties cannot trated in Fig. 4.2. The relatively low cost be achieved in polymeric solutions such as dianhydride, 3,3',4,4'-benzophenonetetracarpolyamide acids. Instead, the solutions boxylic dianhydride (BTDA), is first rapidly become unacceptably viscous when prereacted with ethanol using NMP as the solsolids contents in excess of 15-20% range are vent. Then, if m-phenylenediamine (MPD) is reached thus making them too difficult for added to the solution, Skybond 700 results. If prepregging using commercially available 4,4'-methylenedianiline (MDA) is employed, then Skybond 703 is produced. During cure equipment. In order to prepare polyimide binder solu- the application of heat causes the elimination tions (or polyimide precursor solutions as they of the solvent along with 2 moles of water and are sometimes called) it is necessary to have 2 moles of ethanol per repeat unit to produce the aromatic dianhydride in either one of the the polyimide. The molecular weight initially two possible open ring forms (tetraacid or achieved will depend on the monomer imbaldiester diacid). If the tetraacid form is com- ance employed. It has long been speculated mercially available then the binder solution that during the cure process branching can can be made directly because there is essen- occur by the reaction of amine end-groups tially no reaction with the diamine until with the bridging carbonyl group of the BTDA
Condensation polyimide chemistry 77
3,3',4,4'-benzophenonetetracarboxylic dianhydride BTDA
1
I
m-phenylene diamine
Skybond700
Diethyl ester of 3,3',4,4'-benzophenone tetracarboxylic acid BTDE
4,4'-methylene dianiline
+
Skybond703
Fig. 4.2 Chemistry of SkybondO binder solutions.
moiety leading to branching and intractability of the matrix resin. This is probably one of the reasons why it is difficult to fabricate low void composites using this type of chemistry. 4.2.3 CHEMISTRY OF NR-150
The formation of DuPont's monomeric binder solution NR-150 (used to make Avimid@N prepreg) is illustrated in Fig. 4.3. The 2,2-bis(3',4'-dicarboxyphenyl) hexafluoropropane (6F tetraacid) is dissolved in a suitable solvent (e.g. ethanol, diglyme or NMP) along with a 95/5 mixture of p-phenylenediamine/rn-phenylenediamineto form a low viscosity, stable monomeric solution' suitable for prepregging. During the cure both solvent and 2 moles of by-product water per repeat unit are eliminated to initially form the transient intermediate polyamide acid. According to the studies of Sonnett et aL2, although a low concentration of amide acid persists for a relatively long time period during the early stages of cure, most is rapidly converted to the imide form by the thermally induced elimination of water (2 moles per repeat unit). Since the tetraacid and diamines are very pure, highly
reactive and present in stoicluometric proportions, extremely high molecular weight polyimide is ultimately produced. Also, since the bridging hexafluoroisopropylidene group is inert under normal cure conditions, the polyimide produced is essentially linear and & that sense thermoplastic. 4.2.4 CHEMISTRY OF 3F/36F POLYIMIDES
One of the keys to success in producing an essentially linear condensation polyimide is to have a chemically inert flexibilizing linkage in the dianhydride moiety. As indicated in Section 4.2.3 one of the ways of doing t h s is to employ a hexafluoroisopropylidene bridging group. Since the phenyl group is also very inert another approach is to replace one of the CF, groups with a phenyl group3. This results in the so-called 3F dianhydride. When the 3F monomer is used along with PPD the polyimide produced is called 3F-PPD polyimide. If, on the other hand, a mixture of the 3F and 6F monomers are employed along with the same diamine the copolyimide is designated 36F-PPD.
78 High temperature resins 0
H O -O C
CF3
-
!/D
II
~
F
00 NHZ
D /~
0
- - OOH H + C
/
0
NH2
+
/ NH,
+
Solvent
NH2 956 mixture PPDA4PD
Monomeric binder solution
f
J
Intermediate transient polyamide acid (-2H20)
r
L
1
NR-150 Polyimide
Fig. 4.3 NR-150 polymerization chemistry.
4.3 ADDITION POLYIMIDE CHEMISTRY
4.3.1 OVERALL CHEMISTRY OF ADDITION POLYIMIDE PRECURSOR SOLUTIONS
One of the important concerns in the cure of conventional condensation polyimides is the proper management of the evolution of volatile by-products. During the early stages prior to pressurization the composite is generally somewhat porous and volatiles can readily diffuse out. However, once pressure has been applied and a low void state is achieved the diffusivity dramatically decreases making it much more
difficult for these volatiles to escape. Therefore, ideally one would want to eliminate all of the volatiles prior to pressurization. In order to accomplish this in a conventional condensation polyimide it is necessary to have a high enough monomer imbalance so that the polymer molecular weight and melt viscosity will be low enough to allow for complete consolidation. The problem is that this molecular weight is generally so low that the matrix resin properties are adversely affected. Properties such as strength, toughness and T, will be much lower than desired.
Addition polyimide chemisty 79 A common way of making a low void composite is to produce in situ an imide oligomer having a low enough molecular weight so that it has good melt flow and readily consolidates when pressure is applied and then, with further heating, can go on to produce much higher molecular weight polymer without the further evolution of volatile by-products. This can be accomplished through the use of reactive endgroups which are capable of reacting with one another in various ways without the evolution of volatiles. However, it must be pointed out that such improvements in processibility are only achieved with some sacrifices in other properties. Because of the presence of crosslinks the toughness can be adversely affected. Also, the kinds of linkages produced during these addition polymerization reactions are generally aliphatic in nature thus making them much more prone to thermal-oxidative attack. Many approaches, which are now discussed, have been taken in order to achieve various trade-offs in processibility against properties.
4.3.2 CHEMISTRY OF PMR-15
The first approach to a commercially successful addition polyimide was PMR-15. The basic PMR (Polymerizationof Monomeric Reactants) chemistry was originally invented and developed at NASA Lewis Research CenteF. As illustrated in Fig. 4.4 a monomeric solution is first prepared consisting of the dimethyl ester of BTDA (BTDE), the diamine, MDA and the monomethyl ester of nadic anhydride (NE),the reactive end-capping agent. The monomer ratios are n:n+l:2 respectively. Because the high solids solution is monomeric it is ideally suited to prepregging. As a result of the application of heat, solvent is eliminated along with the water and alcohol of imidization to produce the intermediate imide oligomer having a formulated molecular weight of about 1500. This normally occurs between 121°C (250°F) and 232°C (450°F). At this point essentially all of the volatile by-products have been eliminated. The stage is now set for pressurization. At some-
v
2 Moles NE
2 Moles MDA PMR-15 Binder Solution
Fig. 4.4 PMR-15 polymerization chemistry.
CrosslinkedPMR-15
1 Mole BTDE
80 High temperature resins what higher temperatures the imide oligomer undergoes melting and, if pressure is applied, consolidation to form a low void structure occurs. At temperatures in the 275-316°C (527400°F) range ring opening of the nadic end-groups takes place in a reverse Diels-Alder type of reaction first discovered by Lubowitz9Jo in 1970 and a complex series of reactions takes place involving these end-groups leading to high molecular weight polymer without the further evolution of volatile by-products. Hence the low void state originally achieved during the initial consolidation is maintained and a h g h quality laminate usually results.
4.3.3 CHEMISTRYOF PMR-I1
During the early 1970s it became clear that although composites based on PMR-15 possessed good enough properties to make them a commercial success there was still room for improvement. For instance, the relatively high aliphatic content which came from the nadic end-groups and the methylene group of the MDA contributed to thermal-oxidative instability. Also, the relatively high cross-link density in a polyimide with a formulated molecular weight of 1500 resulted in a rather brittle matrix resin. In order to overcome these deficiencies PMR-I1 was invented”-14(Fig 4.5). For PMR-I1 polyimides the 6F dianhydride is used in place of the BTDA and is converted in situ to the diethyl ester diacid derivative
1
0
PMR-I1 imide prepolymer -l
L
V-CAP imide prepolymer
$$ CH, -CH,
CYCAP imide prepolymer
L AFR700B imide prepolymer
Fig. 4.5 Structures of 6F based addition polyimide prepolymers.
Addition polyimide chemistry 81 (6FDE)by pre-reacting with ethanol. The sin- 4.3.6 CHEMISTRY OF AFR700B gle ring diamine, PPDJis employed in place of With any of the reactive end-groups discussed the MDA. The end-capping agent is still nadic thus far in this chapter the concentration of anhydride. The formulated molecular weight end-capping agent has always been such that was increased from 1500, which is what it is in there are end-caps on both chain ends. PMR-15, to the 3000-5000 range (yields However, by adding just enough end-capping PMR-11-30 and PMR-11-50 respectively). While agent to cap one chain end only and assuming the polymerization is basically the same in that high molecular weight polymer can still both PMR-I1 and PMR-15, since the aliphatic somehow be obtained under reasonable procontent has been greatly reduced, significant cessing conditions, a further significant improvements in thermal-oxidative stability improvement in thermal-oxidative stability and toughness can be realized. should be possible simply because the concentration of unstable groups will have been cut in half. This is exactly what has been done 4.3.4 CHEMISTRY OF V-CAP with AFR700B based on the work of Serafini et Assuming that the 6F/PPD backbone has near al. at TRWz1-23. optimal thermal-oxidative stability and formuIn AFR700B the ratio of 6FDE:PPD:NE is lated molecular weights in the 3000-5000 range 8:9:1 and the formulated molecular weight is are about right from the stand-point of melt about 4400 (see Fig. 4.5 for its s t r ~ c t u r e )This ~~. flow, if further improvements in stability are to means that the nadic end-group component is be realized then one approach would be to only 3.7% of the overall weight compared with make changes in the end-group chemistry. One 22% for PMR-15 and 6.5% for PMR-11-50). effective method is the replacement of nadic During cure of AFR700B composites it is necanhydride as the end-capping agent with p essary not only to achieve the normal aminostyrene The imide prepolymer imidization and nadic end-group coupling that is produced is illustrated in Fig. 4.5. In this reactions but also other unspecified reactions case there is only one weak bond per end-cap involving the amine end-groups undoubtedly compared with eight for the nadic end-cap. occur leading to the formation of the desired During the addition polymerization phase of strong, stiff, tough, high Tgpolyimide. the cure free radical polymerization of the vinyl (-CH=CH,) end-groups occurs leading to high 4.3.7 CHEMISTRY OF TRW-R-8XX molecular weight polymer. The most recent addition to the growing family of commercially available polyimides comes from TRW and has been designated Another approach to a more stable polymer TRW-R-8XXZ5.Although the chemical struchas been to replace the nadic end-capping ture of this polyimide has not yet been agent with 2-amino-p-cyclophane (APC)19,20.disclosed it is reported to be a condensaThe structure of the imide prepolymer con- tion/addition polyimide based on relatively taining the CYCAP end-groups that is initially low cost monomers, making its cost comparaproduced is illustrated in Fig. 4.5. This type of ble to that of PMR-15. It is reported to be free end-capping agent has only two weak bonds of the carcinogen, MDA. per end-cap. Thermolytic cleavage of the -CH,-CH,- groups during the final stages of cure produces bi-radicals (-CH; CH,-) which undergo coupling with other bi-radicals to build up the molecular weight. 4.3.5 CHEMISTRY OF CYCAP
82 High temperature resins 4.4 COMMERCIAL AVAILABILITY OF BINDER SOLUTIONS
Most of the binder solutions described in this chapter are not commercially available but rather are prepared by the prepregger on an as-needed basis just prior to prepregging. One notable exception is Monsanto's Skybond. The monomeric solutions (Skybond 700 and 703) have solids contents in the 45-52% range and solution viscosities of 3000-7000 poisez6. 4.5 COMMERCIAL AVAILABILITY OF PREPREGS
At the time of writing of this chapter all of the different polymide prepreg systems were commercially available from one prepregger or another. However, the reader should appreciate the fact that as time goes by some prepreg types will disappear from the market place and others with an improved balance of processing characteristics, properties and economics will come along to take their place. Also, some companies will go out of business or will be bought out by other companies as consolidation in the industry occurs. Therefore, if a given type of prepreg is required the reader should contact their favorite prepregger. If that particular company does not offer the product the reader requires then advice should be obtained as to where such prepreg could be commercially purchased, if at all. Another fruitful source of information would be to search the Internet.
Table 4.1 Typical properties of neat cured NR-150
Units
Property
''
Value
"C "F g cm-?
350-370 662-700 1.43-1.45 Density oc-I 5.6 x 10-5 Coefficient of thermal OF-' 3.1 x expansion % 60 Char yield MPa (ksi) 110 (16) Tensile strength % 6 Elongation, RT 65 316°C (600°F) J m-2 (in lb in") 2000 (11.4) Fracture toughness Rochwell hardness 70 (E scale)
Table 4.2 Typical properties of neat cured PMR-15
Property Density T, Coefficient of thermal expansion Tensile strength Tensile modulus Elongation Compressive yield strength Compressive strength Equilibrium moisture absorption Fracture toughness
Units
Value
"C (OF) 335 (635) g ~ r n - ~ 1.30-1.32 oc-1 16x1O4 OF-1 28X 10"
MPa (ksi) MPa (ksi)
Reference 28 28 35 35
MPa (ksi)
55 (8.0) 3200 (470) 1.5 110 (16)
35 35 36 35
MPa (ksi)
186 (27)
35
96
4.2
28
"/o
J m-2 500 (2.86) (in lb i r 2 )
37
In neat resin thermal-oxidative stability studies carried out by Sco1az7it was found that 4.6 NEAT RESIN PROPERTIES after 24 h at 316°C (600°F) the NR-150 resin The neat resin mechanical properties for cured had lost 9% of its weight compared with about NR-150 and PMR-15 are summarized in Tables 76%for the PMR-15. Neat AFR700B, which has 4.1 and 4.2 respectively. One of the important been post-cured under nitrogen, has been differences between these two resin systems is reported to have a room temperature tensile their toughness. Cured NR-150 has been strength of 93.8 MPa (13.6 ksi) with 18% retenfound to be dramatically tougher (2000 Jm-' tion of this value at 371°C (700°F). No fracture toughness) compared with a value of properties were available for Skybond, PMR-11, 500 Jm-' for the cured PMR-15. Another signif- V-CAP and CYCAP. Very little information icant difference is thermal-oxidative stability. was available for the 3F-PPD and 36F-PPD
Processing characteristics polyimides except for neat resin densities (1.35 and 1 . 4 2 g ~ m -for ~ the 3F and 36F resins respectively) and Tg of 365-370°C (689-698°F) for both systems in the as-molded state and 405-410°C (761-770°F) for the post-cured state3.TRW-R-8XX has been reported25to yield polyimides having T g in the 400-426°C (750-800°F) range and composite weight loss characteristics at 371°C (700°F) up to 10 times better than PMR-15. All of the resin systems described in this chapter appear to possess good strength and stiffness. Thus, provided that complete cure is achieved during processing and low void composites are produced possessing good fiber/matrix adhesion, high levels of composite mechanical properties should be obtained with good retention (at least 50%) of these properties to just below their Tg. 4.7 PROCESSING CHARACTERISTICS
4.7.1 GENERAL COMMENTS
One of the features that can clearly differentiate one resin system from another is the ease with which a fully cured low void composite can be produced having a specified fiber volume. Also, although all polyimides can have their Tgincreased from post-cure, some systems respond much more readily than others. In all cases there is a problem of properly managing the release of a significant amount of volatiles (normally 10-15% of the weight of the prepreg). The ways in which this can be accomplished can vary significantly from one resin to another depending on the chemistry involved. Factors such as the techniques employed during lay-up of the vacuum bag assembly, pressure, heat-up rate, maximum cure temperature, vacuum application and intermediate holds all must be carefully controlled and optimized for each system. Unfortunately the story on the processing of the various systems covered in this chapter is very incomplete. In a majority of cases information on processing was simply not available
83
because such data was either considered proprietary or classified as secret by the Government Laboratories. In other cases, where some processing information was available in the open literature, very little was usually said about the quality of the part produced by a given cycle so it is difficult to compare the processibility of one resin system with another since the quality of the laminates produced is generally unknown. To make matters worse there are also the issues of the processing characteristics of thick compared with thin sections and how processing can be handled, if at all, when both thin and thick sections are present simultaneously in a given part. It is, therefore, of paramount importance for workers in this field to have as clear an understanding as possible of the chemistry involved at every stage of the cure so that they can quickly and efficiently develop the optimum cure cycle for a given part. The cure cycles presented below will give the reader an approximate idea of the kinds of conditions that have been employed to produce a part. 4.7.2 SKYBOND PROCESSING CONDITIONS
The following autoclave cure cycle has been recommended by MonsantoZ6for 12 ply (3.2 mm, 0.125 in thick) 181 style E-glass (soft A-1100 finish) fabric/Skybond 700 laminates: 0 0
0 0 0 0
apply full vacuum; heat to 177°C (350°F) at 1.7-2.8"C/min (3-5°F /min); hold 5 min at 177°C (350°F); apply 0.69 MPa (100 psi); hold 30 min; cool under pressure and vacuum.
In order to maximize high temperature properties it is recommended that a post-cure should be carried out up to and including the expected use-temperature. For laminates of the type described above a post-cure is suggested in which the part is heated to 200°C (392"F), 225°C (437"F), 250°C (482"F), 300°C (572"F), 325°C (617"F), 350°C (662°F) and
84 High temperattive resins 371°C (700°F) and held for 2 h at each temperature. It is recommended that if thicker laminates are involved the post-cure cycle should be extended. For the same type of laminate based on Skybond 703 a similar autoclave cure cycle can be employed. The only significant differences are a slower heat-up rate (l.l-1.7"C/min, 2-3"F/min) with the pressure being applied at 121°C (250°F)on the way to the final cure temperature of 177°C (350°F).Also a similar postcure cycle is suggested in order to achieve maximum heat resistance. While there is very little definitive information on the void content of Skybond based laminates it is believed that they are generally in the 5-20% range. A most important feature of this particular polyimide system is that the maximum autoclave processing temperature is only 177°C (350°F). No other resin system described in this chapter can make that claim.
4.7.3 PMR-15 PROCESSING CONDITIONS
The following represents a typical autoclave cure cyclez8cited for PMR-15: apply 7-21 kPa (1-3 psi) vacuum; raise autoclave temperature to 227°C (440°F)at 0.83-1.l0C/min (1.5-2.O0F/min); at 163-177°C (325-350°F) apply full vacuum; dwell at 227°C (440°F) for 1 4 h depending on part thickness (up to 2.8 mm, 0.11 in = 1 h, 2.8-6.4mm, 0.11-0.25 in = 2 h, 6.4-12.8 mm (0.25-0.50 in) = 3 h); raise temperature to 238°C (460°F) at 1.1"C/ min (2.0°F/ min); hold at 238°C (460°F) while 1.38MPa (200 psi) autoclave pressure is applied. Do not hold longer than 15 minutes while pressure is being- applied __ raise temperature to 316°C (600°F) at 2.2-3.3"C / min (46°F/min); dwell at 316°C (600°F) for 3 h; cool to room temperature. Vent autoclave pressure below 204°C (400°F) It should be
0
0
0 0
0 0
noted that pressurization does not occur until a temperature of about 238°C (460°F) is reached. At that temperature essentially all of the solvent and imidization volatiles have been eliminated. Normally PMR-15 laminates are subjected to an oven post-cure: heat from room temperature to 204°C (400°F) at 5.6"C/min (10"F/min); heat from 204°C to 288°C (400°Fto 550°F) at 1.1"C/min (2"F/min); dwell at 288°C (550°F)for 1 h; heat from 288°C to 316°C (550°F to 600°F) at 1.1"C/ min (2"F/min); dwell at 316°C (600°F)for 10-16 h; cool to room temperature at 2.8"C/min (5"F/min) maximum.
4.7.4 PMR-I1 AND V-CAP PROCESSING CONDITIONS
The autoclave cure cycles for PMR-I1 and V-CAP based composites are similar to that of PMR-15. The main difference is that the maximum processing temperature has been increased from 316°C to 371°C (600°F to 700°F). The following typical autoclave cure and post-cure cycles for graphite reinforced PMR-I1 composites has been rep~rted'~? 0 0
0
0
0
0
apply full vacuum at room temperature; heat to 149°C (300°F) at 3.9"C/min (7.O"F/ min); hold for 30min at 149"C, then apply 172 kPa (25 psi); heat to 288°C (550°F) at 3.9"C/min (7.O0F/min) with the pressure being increased to 344 kPa (50 psi) at 177°C (350°F) and then to 1.38MPa (200psi) at 232°C (450°F); heat from 232°C to 371°C (450°F to 700°F) at 2.8"C/min (5"F/min); cool under full pressure and vacuum to 232°C (450°F)slowly; cool from 232°C to temperature rapidly.
As with PMR-15, a post-cure is normally carried out in a circulating air oven:
Mechanical properties before and after air aging 85 heat from room temperature to 260°C their high T , (377418"C, 710-785°F) the reten(500°F) at 20"C/min (36"F/min) tion of properties was excellent out to 0 heat from 260°C to 385°C (700°F to 725°F) at temperatures as high as 360°C (680°F). l"C/min (1.8"F/min) with 2 h holds at 316°C At the writing of this chapter the long term (600"F),343°C (650°F)and 20 h hold at 385°C. air aging characteristics of laminates of this particular type had not been completed. However, because of its all aromatic character, 4.7.5 PROCESSING CONDITIONS FOR OTHER ultra-high molecular weight and the complete RESIN SYSTEMS absence of any aliphatic character from reacAt the time of the writing of this chapter no tive end-capping agents it should air age well unclassified processing information was avail- and the matrix in Avimid N-150 has been preable for CYCAP, AFR700B, TRW-R-8XX or the viously shown to possess outstanding 3F/36F polyimides. The compression molding thermal-oxidative stability30. conditions used by DuPont to make the laminates whose properties are described in 4.8.3 3F/36F POLYIMIDES Table 4.4 were not disclosed. However, details concerning the autoclave processing of Although this family of all aromatic polyAvimid N have been previously discussedz9. imides is relatively new, preliminary data indicates that high quality laminates possessing good mechanical properties and excellent 4.8 MECHANICAL PROPERTIES BEFORE AND long term thermal-oxidative stability can be AFTER AIR AGING produced. According to the work of Scola3 both the 3F-PPD and the 36F-PPD systems 4.8.1 SKYBOND resulted in G40-600 laminates having the expected good room temperature mechanical The most common type of reinforcement that properties (flex strength and short beam shear has been employed with Skybond binders is strength) with excellent retention of these 'dry' E-glass fabric. The mechanical properties of properties out to at least 371°C (700°F). Also, as-molded and air aged laminates based on after 100 h in air at 371°C (700°F) these lamiSkybond 700 are summarized in Table 4.3. This nates retained at least 100% of these properties particular resin system has been tailored for with weight loss values ranging from 1.4%for extended exposures at temperatures up to the 36F-PPD resin and 2.4% for the 3F-PPD 371°C (700°F). For 343°C (650°F) applications polyimide. Surprisingly neat resin studies carSkybond 703 is recommended by the manuried out by Scola have similarly shown that facturer. It is interesting to note that in spite of the 36F-PPD polyimide appeared to be somethe relatively high porosity levels (5-20%) what more stable than the 3F-PPD polymer. Skybond binders are still in demand for cerFor instance, after 100 h exposure at 371°C tain specialty applications. (700"F), the 36F-PPD copolyimide had lost about 2% of its weight compared with 3.2% for the 3F-PPD polymer. Much more work needs 4.8.2 AVIMID N to be carried out on polyimides based on the Some mechanical properties for compression 3F monomer before any final decision can be molded/post-cured Celion G30-500 miweave/ made as to its long term viability in the market Avimid N laminates are tabulated in Table 4.4. place and how it will ultimately compete with All laminates had a very low void content the 6F based polyimides such as Avimid N, (4'30)and, therefore, high levels of mechanical PMR-11, V-CAP, CYCAP and AFR700B. properties at room temperature. Because of
86
High temperature resins
Table 4.3 Mechanical properties of Skybond@700/181 style E-glass laminates (Soft A-1 100 Finish) 26
High temperature, high pressure
Vacuum bag
Flex strength, MPa (ksi) 24 "C (75 OF)
517-586 (75-85)
524-576 (76-84)
371°C (700°F)after 0.5 h at 371°C 371°C (700°F) after 100 h at 371°C
310414 (45-60) 138-241 (20-35)
152-221 (22-32) 138-166 (20-24)
316°C (600°F) after 500 h at 316°C 316°C (600°F)after 860 h at 316°C 316°C (600°F) after 1850 h at 316°C
200 (29) 138 (20) 76 (11)
288°C (550°F)after 2300 h at 288°C 288°C (550°F)after 4500 h at 288°C 288°C (550°F) after 9000 h at 288°C
283 (41) 221 (32) 103 (15)
Flex modulus, GPa (msi) 24°C (75°F)
22 (3.1)
19 (2.8)
299°C (570°F) after 335 h at 299°C
22 (3.1)
-
371°C (700°F) after 100 h at 371°C
-
12 (1.8)
316°C (600°F) after 500 h at 316°C 316°C (600°F) after 860 h at 316°C 316°C (600°F)after1850 h at 316°C
18 (2.6) 18 (2.6) 14 (2.1)
288°C (550°F) after 2300 h at 288°C 288°C (550°F) after 4500 h at 288°C 288°C (550°F)after 9000 h at 288°C
18 (2.6) 20 (3.0) 14 (2.0)
Property
Ultimate tensile strength, MPa (ksi) 24°C (75°F) 299°C (570°F)after 335 h at 299°C 24°C (75°F) after 100 h at 250°C (482°F) 24°C (75°F) after 100 h at 300°C (572°F) Elongation, YO 24°C (75°F) 299°C (570°F) after 335 h at 299°C 24°C (75°F) after 100 h at 250°C (482°F) 24°C (75°F) after 100 h at 300°C (572°F) Weight loss, YO After 100 h at 371°C (700°F) After 500 h at 316°C (600°F) After 860 h at 316°C (600°F) After 1850 h at 316°C (600°F) After 2300 h at 288°C (550°F) After 4500 h at 288°C (550°F) After 9000 h at 288°C (550°F)
393 (57) 290 (42)
347 (50.1)
-
336 (49) 332 (48.2)
-
-
1.9 1.4
2.0
-
1.7 2.0
-
3.0 2.2 3.4 7.9 3.6 5.0 12.0
-
Mechanical properties b$ore and after air aging 87 Table 4.4 Mechanicalproperties of compression molded Celion@G30-500 Uniweave/Avimid@N laminates 39 No. ofplies
Orientation
Tf "C (OF)
Temperature, "C
Property
(OF)
16
0"
418 (785)
24 (75) 218 (425) 316 (600) 360 (680)
Short beam shear strength, MPa (ksi) 98.6 (14.3) 61.4 (8.9) 46.2 (6.7) 38.6 (5.6)
16
0"
418 (785)
24 (75) 316 (600) 360 (680)
Flex strength, MPa (ksi) 1344 (195) 731 (106) 565 (82)
16
0"
418 (785)
24 (75) 316 (600) 360 (680)
16
(k45")
418 (785)
24 (75) 371 (700)
10
0"
377 (710)
24 (75) 416 (780)
10
0"
377 (710)
24 (75) 416 (780)
10
0"
377 (710)
24 (75) 416 (780)
10
0"
377 (710)
24 (75) 416 (780)
16
0,90,i45"
377 (710)
24 (75) 360 (680) 416 (780)
16
0,90,+45"
377 (710)
24 (75) 360 (680) 416 (780)
16
0,90,45"
377 (710)
24 (75) 360 (680) 416 (780)
16
0,90,i45"
377 (710)
24 (75) 360 (680) 416 (780)
16
0,90,i45"
377 (710)
24 (75)
16
0,90,+45"
377 (710)
24 (75) 360 (680) 416 (780)
16
0"
377 (710)
24 (75) 316 (600) 360 (680)
16
0"
418 (785)
24 (75) 316 (600) 343 (650)
Flex modulus, GPa (msi) 126 (18.0) 123 (17.9) 117 (17.0) In-plane shear strength, MPa ( h i ) 174 (25.2) 104 (15.1) Tensile strength, MPa (ksi) 1261 (183) 1027 (149) Tensile modulus, GPa ( m i ) 124 (18.0) 104 (15.1) Open hole tensile strength, MPa (ksi) 1027 (153) 854 (124) Open hole tensile modulus, GPa (msi) 132 (19.1) 114 (16.5) Tensile strength, MPa (ksi) 460 (66.7) 314 (45.5) 236 (34.3) Tensile modulus, GPa (msi) 56.5 (8.2) 46.8 (6.8) 30.3 (4.4) Open hole tensile strength, MPa (ksi) 389 (56.5) 250 (36.3) 177 (25.5) Open hole tensile modulus, GPa (msi) 54.4 (7.9) 46.8 (6.8) 38.3 (5.7) Compressive strength, MPa (ksi) 389 (56.5) Compressive strength, MPa ( h i ) 458 (66.4) 312 (45.4) 270 (39.2) Compressive strength, MPa (ksi) 868 (126) 448 (65.1) 409 (59.4) Interlaminar fracture toughness J m-2 (in Ib in-') 630 (3.6) 682 (3.9) 718 (4.1)
88 High temperature resins Table 4.5 Mechanical properties of fiber reinforced PMR-15 laminates 35
Reinforcement Property
Compressive strength MPa (ksi) 23°C (73°F) 288°C (550°F) Compressive modulus GPa (msi) 23°C (73°F) 288°C (550°F)
High strength 1>3447 MPa (500 ksi)] Standard modulus /228 GPa (33 msi)] class Carbon fiber unidirectional tape (57-63 Vol. %fibers)
High strength 1>3447 MPa (500 ksi)] Standard modulus 1228 GPa (33 msi)] class Carbon fiber 8-harness satin fabric (55-60 Vol. %fibers)
517-586 (75-85)
827-965 (120-140) 758-896 (110-130)
552-689 (80-100) 414-552 (60-80)
28-34 (4-5)
97-117 (14-17) 83-110 (12-16)
62-76 (9-11) 48-62 (7-9)
7781 Style E-glass fabric (50-55 Vol. %fibers)
Flex strength MPa (ksi) 23°C (73°F) 288°C (550°F) 316°C (600°F)
483421 (70-90)
Flex modulus GPa (msi) 23°C (73°F) 288°C (550°F) 316°C (600°F)
21-34 (3-5)
965-1103 (140-160) 689-896 (100-130)
414-552(60-80)
5 5 4 9 (8-10) 5 5 4 9 (8-10)
21-34 (3-5)
Tensile strength MPa (ksi) 23°C (73°F) 288°C (550°F)
1241-1448 (180-210) 1241-1448 (180-210)
689-896 (100-130) 758-965 (110-140)
Tensile modulus GPa (msi) 23°C (73°F) 288°C (550°F)
117-138 (17-20) 103-124 (15-18)
62-76 (9-11) 62-76 (9-11)
Interlaminar shear strength MPa (ksi) 23°C (73°F) 288°C (550°F) 316°C (600°F)
62-76 (9-11) 34-48 (5-7)
5 5 4 9 (8-10) 34-48 (5-7)
Mechanical properties before and after air aging 89 4.8.4 PMR-15
perature and either 232°C (450°F) or 288°C From the data cited in Table 4.5 it can be seen (550°F) at heat-up and cool-down rates of that PMR-15 based composites reinforced with 278°C (532"F)/min (achieved by employing a either E-glass or graphite fibers are character- heated fluidized sand bed) significant microcized by good room temperature strength racking was detected at 5000 thermal cycles properties with excellent retention of these and beyond. The data plotted in Fig. 4.6 properties at temperatures in the 288-316°C clearly indicate that cycling to the higher temperature resulted in a much higher (550-600°F) range. concentration of microcracks.According to the Over the past 20 years a great deal of effort results plotted in Fig. 4.7 the development of has gone into the study of the weight loss microcracks was found to have a very deletecharacteristics of PMR-15 composites at varirious effect on the matrix-dominated ous temperatures and air pressures and the compressive strength of the graphite fabric effects of these exposure conditions on the laminates. However, GE also showed that the laminate mechanical properties. One of the fiber-dominated tensile strength was not significant new property features to be uncovaffected by the presence of the microcracks. ered has been the effects of temperature There is every reason to believe that composcycling on the development of microcracks ites based on other resin systems will also and the effects of these microcracks on the undergo microcracking from temperature strength and stiffness properties. In accelerated thermal cycling studies3I carried out at cycling and can be expected to exhibit General Electric - Aircraft Engines in which improved microcracking performance to the the cycling was carried out between room tem- extent to which they possess improved toughness compared with PMR-15. 35
c
I
A
A A
15
1
A A
A
.-c
u)
n
- 288 "C (550 'F) RT - 232 "C (450 OF) RT
Y
u 0 2
5
Data
Data
1 -
0 0
5000
10000
15000
20000
25000
Number of Cycles
Fig. 4.6 Crack density compared with accelerated thermal cycles for graphite fabric/PMR-15 laminates. 31
90 High temperature resins 100 1
I
1
I
I
0
P 0
P P
A A A
-
RT 288 "0(550
A t
O A 0
OF)
Data
A I
I
I
5000
10000
15000
I
I
20000
25000
Number of Thermal Cycles Fig. 4.7 Effect of accelerated thermal cycling on the compressive strength of PMR-l5/graphite fabric laminate~.~~ 4.8.5 PMR-I1
The greatly improved weight loss performance of PMR-I1 based composites compared with PMR-15 as a result of aging in air at elevated temperatures (343"C, 650°F) is illustrated in Fig. 4.814,15. Note that very little difference in weight loss was found between PMR-11-30 and PMR-11-50. Although a stability improvement would have been expected with the higher formulated molecular weight it usually is not detected because of the difficulty in making low void laminates with the higher viscosity resin. Porosity itself, of course, can contribute to thermal-oxidative instability. As a general rule, for any of the polyimides discussed in this chapter, the high temperature mechanical properties tend to increase as a result of the air aging process. This is presumably because the oxidative crosslinking that occurs during the aging process tends to increase the T,. In Fig. 4.9I4J5 it can be seen that
the 343°C (650°F) interlaminar shear strength increased for the first 200-400 h of aging. Even after 600 hours aging the interlaminar shear strengths were about the same as they were at the beginning. However, the reader must be aware that there has been a considerable weight loss (about 8%) with the resulting shrinkage and induced stresses. Also, the matrix resin has presumably become embrittled. Therefore, in assessing the useful lifetime of a composite at elevated temperatures it is important to not only consider the changes in various strength and stiffness properties but also the effects that weight loss itself can have on dimensional stability and matrix resin toughness. Carbon fiber selection can also be an important factor in determining the overall high temperature performance of a composite. As illustrated in Fig. 4.1032PMR-11-50 laminates reinforced with Celion@G40-600 or ThomeP 650-35 retained a higher level of flex strength
Mechanical properties before and after air aging 91 I
I
12 10
0
A
PMR-15
0
PMR-11-30
0
PMR-i1-5(i
100
200
I
I
I
I
400
500
600
700
I
I 300
Time, Hours
Fig. 4.8 Weight loss of Celiona 6KJPMR laminates at 343°C (65O0F).l4Js
I
60
I
I
I
0
I
I
1
0
.y"
OA
40
0
?!
3i
I 0
8.70
7.25
0
G
5
r;p
I
' 0
I
I
I
I
I
I
100
200
300
400
500
600
4.35
!
700
Time, H ow
Fig. 4.9 Interlaminar shear strength of Celiona 6K/PMR laminates after aging in air at 343°C (650°F).14,15
92 High temperature resins 180
2 00
160
160
.-
u)
Q
k!
s' 0
-
Y
120
120
tl, C
?!?
z
C
!
cy)
5-
80
*O
40
40
X
8
x
@ u,
ii
0
0
After 500 Hours at 371 OC (700 O F )
Initial
Fig. 4.10 Effect of fiber selection on the flex strength of PMR-11-50 laminates air aged at 371°C (700°F).32
100
rI
1I
I
1
800 Hours' Exposure
T40R
I
G4U-600 G40-700 G40-800 T650-42 T650-35 Fiber
Fig. 4.11 Thermal-oxidative stability of various graphite fibers exposed to 371°C (700 O F ) air. 32
Mechanical properties before and after air aging 93 after 500 h at 371°C (700°F)than laminates reinforced with other fibers such as Thornel 40R, Thornel 650-42, Celion G40-700 and Celion G40-800. It is of interest to note that according to the data in Fig. 4.1132there was not a one-toone correlation with the basic carbon fiber stability. While Celion G40-600 had one of the lowest weight losses in 371°C (700°F) air aging the Thornel 650-35 had one of the highest. 4.8.6 V-CAP AND CYCAP
performance in comparison with PMR-11. These same researchers have also shown that V-CAP/graphite compositespossess improved thermal-oxidative stability compared with PMR-11-50. However, with the advent of AFR700B where significant improvements in thermal-oxidative stability have been realized by simply capping one end only with the low cost nadic anhydride, it is not clear whether either the V-CAP or CYCAP polyimides with their more expensive end-capping agents will be successful in the market place.
As a result of the reduced aliphatic character of the reactive end-groups in CYCAP, Meador et a l l 9 have reported improved weight loss Table 4.6 Mechanical properties of unidirectional AFR700B/S2 glass tape laminates 24
Material condit ion
Fiber or ien f af ion
Test temperature, "C ( O F )
VOl. % fibers
23
(73)
316
(600)
371 (700)
Flex strength, MPa (ksi) As-molded Air aged 100 h at 371°C (700°F) (1atm)
0" 0"
55 55
1123 (163) 958 (139)
634
(92)
496 517
(70) (75)
90" Flex strength, MPa (ksi) As-molded Air aged 100 h at 371°C (700°F) (1 atm)
0" 0"
48 (6.9) 23 (3.3)
55 55
29 13
27 (3.9) 23 (1.7)
(4.2) (1.9)
Flex strength, MPa (ksi) ~~~
Cycled 100 x from RT to 371°C (700°F) in air 50 h total at 371°C
0"
55
972 (141)
421 (61)
In-plane shear strength, MPa (ksi) ~~~
As-molded Air aged 100 h at 371°C (700°F) (1 atm)
k 45 f 45
59 59
81 (11.8) 29 (4.2)
53 21
(7.7) (3.0)
41 (5.9) 26 (3.7)
Tensile strength, MPa (ksi) As-molded Air aged 100 h at 371°C (700°F) (1atm)
0" 0"
54 55
1103 (160) 814 (118)
917 703
(133) (102)
834 (121) 676 (98)
0" Compression strength*, MPa (ksi) .~
As-molded Air aged 100 h at 371°C (700°F) (1atm)
0" 0"
58 58
662 621
(96) (90)
386 352
(56) (51)
303 (44) 331 (48)
* Non-standard test; open hole compression fixture used with no-hole specimen. For comparison the room temperature IITRI compression strength (D3410) was 1152 MPa (167 ksi)
94 High temperature resins 2 1
I
I
I
1000
2000
3000
I
I
I
I
II
4000
5000
6000
7000
8000
1.6
% 0
1.2
E
.-0
2"
0.8
0.4
0 0
Time, Hours Fig. 4.12 Effect of air aging at 260°C (500°F) on the weight loss of AFR700B composites.24 4.8.7 AFR700B
One of the most promising new resin systems that has come along in the last several years is AFR700B. One indication of this is the excellent properties tabulated in Table 4.624which had been obtained on unidirectional S2 (933 finish) glass tape laminates, both asmolded as well as after 100 h in air at 371°C (700°F). In long term isothermal air aging studies24 carried out at 260°C (500°F) the weight loss performance of AFR700B laminates reinforced with AstroquartzO I11 fibers was superior to that exhibited by either S2-glass or Thornel 650-42 laminates (Fig. 4.12). The retention of the 260°C (500°F)flex strength of S2-glass laminates is illustrated in Fig. 4.13. Similar results were obtained with Astroquartz I11 and Thornel 650-42. Overall, composites based on AFR700B offer the promise of reasonable processability, cost and excellent mechanical properties in the dry-as-molded state out to about 371°C (700°F). Based on its chemistry
one can also expect significant improvements in both matrix resin toughness and thermaloxidative stability compared with PMR-15. 4.8.8 TRW-R-8XX
Although very little unclassified property data was available at the time of the writing of this chapter, according to TRWZ5,the room temperature mechanical properties of composites based on TRW-R-8XX have been found to be equivalent to those based on PMR-15 but with superior toughness, higher Tg (371°C to > 449"C, 700°F to >840"F), superior retention (>70% compared with <20% for PMR-15) of properties at 371°C (700 OF) and up to 10 times better weight loss characteristics after 100 h at 371°C (700 O F ) compared with PMR-15. Based on these results and assuming reasonable processibility, TRW-R-8XX could turn out to be a serious contender in the world of high temperature polyimides, especially in view of the projected low cost.
Electrical properties 95 90
80 I?
I 70
€
UJ C
?! !
3
v i
60
n t
I
I
I
I
1
I
90
80
.-ua
Y
70
UJ
E \
C
!
3 6o
50
40
40
1000
x
8
ii
50
0
i
2000
3000
4000
5000
6000
7000
Time, Hours Fig. 4.13 Effect of air aging at 260°C (500°F) on the flex strength of AFR700B/S2-glass laminates." 4.9 ELECTRICAL PROPERTIES
One of the useful properties of aromatic polyimides is their good all around electrical properties. Low dielectric constants and dissipation factors have been measured in quartz fabric reinforced Avimid N and Skybond based composite^^^. For instance, the room temperature dielectric constants for Avimid N and Skybond Astroquartz laminates have been
reported to be 3.2 and 3.7 respectively.Also, in the same types of laminates the dissipation factors have been measured to be 0.001 for Avimid N and 0.015 for the Skybond. The partial fluorocarbon character of the 6F monomer is undoubtedly the reason behind the improved properties of Avimid N compared with Skybond. Some other miscellaneous electrical properties26 of E-glass reinforced Skybond based composites are tabulated in
Table 4.7 Electrical properties of Skybond@700/181 style E-glass fabric laminates 26
Property Dielectric strength short time parallel to laminate step by step parallel to laminate short time stepwise Dielectric constant (1MC) Dissipation factor (1MC) Insulation resistance Volume resistivity Surface resistivity
Units
Value
volt volt volts/pm volts/pm
55 000 38 000 179 140 4.1 0.0045 1.9 x 107 2.47 x 1015 3.35 x 1014
megohm ohm-cm Ohm
96 High temperature resins Table 4.7. Skybond 700/E-glass fabric lami- there were also definite indications that voids nates have been shown to have a dielectric in the composite could serve to facilitate the constant and dissipation factor measured at drying out process, thus resulting in a higher X-band (8.5 KMC) frequency at room temper- apparent 'wet' T . This means that although ature of 3.74 and 0.016 respectively. When voids are well known to adversely affect measured at 300°C (572°F) there was essen- mechanical properties their presence could conceivably result in significant improvetially no change. ments in a composite's hot/wet properties to the point where an overall better balance of 4.10 HYGROTHERMAL PROPERTIES properties might be possible. In the design of high temperature polyAll of the mechanical properties discussed thus far in this chapter, either before or after imide parts it is strongly suggested that air aging, have been determined on 'dry' spec- moisture effects be fully taken into account. It imens. Although there are many reports in the seems apparent that the full potential of polyliterature on the absorption of water by a wide imide composites will not be realized until variety of polymers and the effects that this effective ways are found to reduce the adverse water can have on the Tg,strangely there have effects of moisture at elevated temperatures been relatively few reports of the effects of without seriously affecting the other impormoisture on polyimide composites. Hot/wet tant properties such as strength, toughness properties are normally reported for epoxy and thermal-oxidative stability. and bismaleimide composites, but usually not those based on polyimides. Unfortunately for 4.11 END-USE APPLICATIONS polyimide composites water absorption is to be expected since the equilibrium water con- In spite of their recognized limitations (e.g. tent of the neat resins being normally in the microcracking and hygrothermal problems) polyimide composites have been successfully 2 4 % range. One of first references to moisture effects employed i i ~a wide variety of applications. occurred in 1976%in which studies at DuPont For instance, autoclave molded PMR-15 on Avimid N composites indicated that low graphite fabric composite has been employed void (
End-use applications 97
Fig. 4.15 AvimidO N/graphite jet engine variable
stator vane bushings. engine as well as various splitters and fairings for the F-110 engine. Compression molded Avimid N/graphite variable stator vane bushings (Fig. 4.15) and washers, now available from DuPont/Tribon Composites, have been extensively employed since the early 1980s in conjunction with the
variable stator vanes in a variety of military and commercial jet engines. Parts of this type have also been made using PMR-15 and PMR-11. Another interesting non-aerospace application for Avimid N from DuPont has been the product, NO-CHX@takeout jaws (Fig. 4.16), which have been used in the glass bottle industry to grab the hot blow-molded glass bottles to transport them during fabrication without cracking or checking. Other applications for polyimide composites have been radomes, missile fins, jet engine nozzle flaps, fairings, cowls and inlet guide vanes, gear cases for helicopters and heat shields.
1
Fig. 4.16 NO-CHX@Take out jaws based on Avimid N/graphite.
98 High temperature resins REFERENCES
1. Gibbs, H.H. 20th Natl. SAMPE Symp. Exhib., April 1975. 2. Sonnett, J.M., McCullough, R.L., Beeler, A.J. and Gannett, T.P. 24th Intern SAMPE Tech. Conf., October 1992, p. T735. 3. Scola, D.A. United Technologies, private communication. 4. Serafini, T.T., Delvigs, P. and Lightsey, G.R. 1. App. Polym. Sci., 1972,16, 905. 5. Serafini, T.T., Delvigs, P.and Lightsey, G.R. US Patent 3 745 149 (July 1973). 6. Serafini, T.T. and Delvigs, P. Appl. Polym. Symp., 1973, 89, (22). 7. Serafini, T.T. Proc. 1975 Intern. Conf. Composite Materials, AIME, New York, 1976,1,202. 8. Serafini, T.T., Delvigs, P. and Lightsey, G.R. NASA TN D-6877 (1972). 9. Burns, E.A., Lubowitz, H.R. and Jones, J.F. NASA CR-72460 (1968). 10. Lubowitz, H.R. US Patent 3 528 950 (1970). 11. Serafini, T.T., Vannucci, R.D. and Alston, W.B. NASA TMX-71984 (1976). 12. Vannucci, R.D. 32nd Intern. SAMPE Symp. Exhib., April 1987. 13. Vannucci, R.D. and Cifani, D. NASA TM-100923 (1988). 14. Vannucci, R.D. and Cifani, D. 20th Intern. SAMPE Tech. Conf., Sept. 1988. 15. Serafini, T.T., Delvigs, P. and Vannucci, R.D. 36th Ann. Tech. Conf. SPI Reinforced Plastics/Composites Inst., Feb. 1981. 16. Vannucci, R.D., Malarik, D.C., Papadapoulos D.S. and Waters J.F. NASA TM 103233. 17. Vannucci, R.D., Malarik, D.C., Papadapoulos, D.S. and Waters, J.F. 22nd Intern. SAMPE Tech. Conf., Nov. 1990. 18. Malarik, D.C. and Vannucci, R.D. NASA CP10039,15-1 (1989).
19. Meador, M.A., Cavano, P.J. and Malarik, D.C. Proc. Sixth Ann. ASM/ESD Adv. Comp. Conf., Oct. (1990). 20. Sutter, J.K. et al., NASA CP-10039,12-1 (1989). 21. Serafini, T.T. et al., US Patent 5 091 505 (Feb. 1992). 22. Serafini, T.T. et al., US Patent 5 149 760 (Sept. 1992). 23. Serafini, T.T. et al., US Patent 5 149 772 (Sept. 1992). 24. Johnson, K.M. Air Force Materials Laboratory, private communication. 25. Serafini, T.T. TRW-Redondo Beach, private communication. 26. Monsanto trade literature. 27. Scola, D.A. 34th Intern. SAMPE Symp., p. 246 (1989). 28. Hexcel trade literature. 29. Gibbs, H.H. 10th National SAMPE Tech. Conf, 1978, p. 21. 30. Gibbs, H.H. 1. Appl. Poly. Sci., Appl. Poly. Symp., 1979, 35,207. 31. Ward, D. General Electric -Aircraft Engines, private communication. 32. Meador, M.A. NASA Lewis Research Center, private communication. 33. Tanikella, M.S. Mosco, J.A. and Rafalski, T.J., 24th Intern. SAMPE Tech. Conf., 1992, p. 687. 34. Gibbs, H.H. 21st Natl. SAMPE Symp. Exhib., 1976, p. 607. 35. Fiberite trade literature. 36. Engineered Materials Handbook - Composites, 1987,1, 79. 37. Wilson, D. High Performance Polymers, 1991,3(2). 38. DuPont trade literature. 39. Iuliano-Picho, D. DuPont Composites, private communication.
SPECIALITY MATRIX RESINS
5
David A.Shimp
5.1 INTRODUCTION
Bismaleimide (BMI) and cyanate ester (CE) resins were first commercialized in the 1970s as 250-300°C Tgclass laminating resins for circuit board substrates. In the early 1980s structural prepregs were introduced to an aircraft industry searching for primary structure composites with higher service temperatures and improved damage tolerance relative to multifunctional epoxy based composites. Both BMI and CE resins have since evolved as easy-toprocess thermosetting resins qualified for 177°C (350°F) hot-wet service. Toughening technologies provide compression-after-impact ratings approaching or matching the damage tolerance of thermoplastic resin composites. Bismaleimides, with higher modulus values and established higher thermal ratings, earned a strong position in military aircraft primary structures with recent selection for the F-22 fighter. Cyanate esters, with superior dielectric loss properties and lower moisture absorption, are strong contenders for radomes, skins covering phase-array antennae, advanced Stealth composites and space structures. 5.2 RESIN CHEMISTRY
Cyanate ester monomers are prepared by reacting bisphenols or polyphenols with cyanogen chloride in the presence of an organic base (Rottloff, 1977).Crystalline monomers are thermally advanced to amorphous prepolymer Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
resins by resin suppliers accomplishing from 1550% of the cyclotrimerization curing reaction using closely controlled reactor processing. Figure 5.1 illustrates formation of the s-triazine ring (cyanurate trimer) by the cycloaddition of three cyanate groups. Formulators catalyze amorphous monomers or prepolymer resins with latent catalysts which promote full conversion to the thermoset polycyanurate with subsequent heating to 177-250°C. Cyanates also serve to cure epoxy resins, forming costeffective hybrids retaining an anomalously high fraction of CE homopolymer properties (Shimp, 1992). Bismaleimide monomers are prepared by the reaction of aromatic diamines with maleic anhydride in the presence of dehydrating agents (Stenzenberger, 1990). Homopolymers of BMI monomers are excessively brittle and in practice are co-reacted with chain-extending diamines, diallyl bisphenols or dipropenyl phenoxides to develop toughness via reduced cross link density. Figure 5.2 illustrates chain extension with aromatic diamine (Bargain, 1971) to form longer linear segments which ultimately crosslink by homopolymerization of maleic double bonds. Kerimid@resins and Fiberite PI molding compounds are examples of commercial BMIs using aromatic amine modification. Figure 5.3 depicts a series of reactions whereby o-allyl-phenols add across the maleic double bond via the 'ene' reaction and a second maleimide enters into a Diels-Alder ring-forming reaction with the now conjugated propenyl residual double bond. The proposed reaction
100 Speciality matrix resins
0
0
I VI
MICHAEL ADDITION
I VI
HOMOPOLYMERIZATION Prepolymer resin
0
0
0
0
Curing via cyclotrimerization
Y
X = alkylidene
Y
Y = alkyl or H
Fig. 5.2 Sequence of chemical reactions for advancing BMI monomers with aromatic diamines to resin adducts by chain extension and ultimately to toughened thermoset plastics. 0 R
i
Thermoset plastic (polycyanurate)
Fig. 5.1 Dicyanates cure by forming triazine rings on heating, advancing to prepolymers (up to 50% conversion) and to thermoset plastics at -O-C=N conversions >60-64%.
mechanism includes isomerization to form the aromatic ring. Crossllnking can occur by a continuation of these reactions (difunctional components) or via residual maleic double bond homopolymerization. The chemistry by which o,o'-diallyl bisphenol A coreacts with and toughens 4,4'-bismaleimido-diphenylmethane (BMI-DAB) is described by Zahir (1978) and
Matrixformulation King (1984). Allyl functional phenoxy compounds follow the same reaction path while propenyl functional phenoxides eliminate the 'ene' reaction step (Stenzenberger,1990).
r!
Ri
@OH
I
+ I
I bv2b
"ENE" REACTION
Ri
OH
1 yy Ri
DIELS-ALDER REACTION
% db
I
HOMOPOLY-
e . 9 MERIZATION OH AROMATIZATION
1I
101
5.3 COMMERCIAL RESINS
The chemical structure of seven commercial di(po1y)cyanate ester monomers is shown in Table 5.1 along with supplier information, physical state and key homopolymer properties. The three crystalline monomers are usually supplied only as amorphous prepolymers in semisolid, hard resin, or ketone laminating solution form. CE homopolymer properties are not affected by prepolymer advancement, which is only an interruption of the ring-forming curing reaction to alter physical state and rheological properties. Monomer asymmetry, e.g. AroCy@L-10, can yield low RT viscosity. CEs have a low toxicity profile and storage stability comparable to epoxies. Table 5.2 describes several BMI monomers, advanced resins, reactive tougheners with allyl and propenyl functionality and RTM resins. Most BMI monomers have a crystalline physical state. Eutectic blends of monomers are available as resolidified melts of lower melt point. Resins prepared by prereacting a molar excess of BMI with tougheners have a powder or hard resin physical state. Allyl and propenyl functional reactive tougheners are usually viscous liquids which serve to dissolve crystalline BMIs at temperatures below 110°C to offer convenient melt processibility.
5.4 MATRIX FORMULATION
-
HoMoPoLYMERIZATION
Fig. 5.3 Sequence Of reactions between Phenols and BMI monomers involves grafting via the 'ene' reaction and fused ring formation via Diels-Alder. Crosslinking with di(po1y) functional components involves completion of these reactions and/or maleic double bond homopolymerization.
5.4.1 MEETING RHEOLOGICAL REQUIREMENTS
Both resin classes offer a wide selection of monomers and prepolymers enabling the fomulator to satisfy the rheological properties of fluid RTM compounds, tack and cohesive integrity of compliant prepreg and the short flow of compression molding compounds. Figure 5.4 illustrates the limiting direct relationship in families of thermosetting resins between fluid monomer (150 mPa s viscosity) On curing. High temperature and Tg temperature polyimides locate Off-sCalein the
102 Speciality matrix resins Trade name/ supplier/ physical state
Polycyanate monomer structure/ arecursor
T.2
Homopolymer property wt.% Dk G, H,O
MHz
J m-2
289
2.5
2.91
140
252
1.4
2.75
175
270
1.8
2.66
140
258
2.4
2.98
190
192
0.7
2.64
210
3.8
3.08
60
1.4
2.80
125
O
e
~~
AroCy B
e c - ~ c ~ ~ + c ECibaNSpecialty Chem.* BT-2000 Mitsubishi GC
CH3
Bisphenol A
Crystal 943
N E c - o o { *)N
p
AroCy M Ciba Specialty Chem.
3
Crystal
$
dH3
CH3
Tetramethylbisphenol F AroCy F Ciba Specialty Chem.
,oOcP*ZN CF3
Crystal Hexafluorobisphenol A AroCy L-10 Ciba Specialty Chem. Liquid Bisphenol E XU-366 Ciba Specialty Chem. Semisolid Bisphenol M Primaset PT Lonza, Inc.
270 to
XU-371 Ciba Specialty Chem. Semisolid
Novolac resin
*CZN
Dicyclopentadienyl bisphenol
XU-71787 Dow Chemical
>350
244
Semisolid
~
* The complete name of the Ciba company supplying AroCy cyanate resins is Ciba Specialty Chemicals Corp., Performance Polymers Group.
Matrix formulation 103 Table 5.2 Commercial BMI monomers, resins (adducts), reactive tougheners and compounds Supplier
Trade name/structure
Description
~
Basic BMI monomer Crystalline powder m.p. 150-160°C
Ciba 0
0
Matrimidm5292A o,o'-DiallylBisphenol A Reactive toughener 12 000-20 000 mPa s at 25°C
Ciba Matrimid 5292B Inspec
CornpimidemMDAB
Basic BMI monomer
Inspec
Compimide 353 Eutectic monomer blend
Resolidified BMI melt 400-1400 mPa s at 110°C
Inspec
Compimide 796 Proprietary BMI adduct
Resolidified BMI resin melt 10004500 mPa s at 110°C
Compimide TM 121
Bisallyl polyphenoxide Reactive toughener 120-250 mPa s at 71°C
Inspec
a
r'
Inspec
Bispropenyl phenoxy benzophenone Reactive toughener 1000-1600 mPa s at 71°C
0
Compimide TM 123 Inspec
Compimide 15 MRK
Powder for molding compounds
Inspec
Compimide 65 FRW
Resolidified melt for filament winding and RTM
Ciba designates Ciba Specialty Chemicals Corp., Performance Polymers. Inspec designates Inspec Fine Chemicals Co.
Table 5.1 (on facing page) Commercial cyanate ester monomers, suppliers, physical states and homopolymer properties. Water absorption is wt.% at saturation. D, = dielectric constant
104 Speciality matrix resins 400
I
300 0
BM'
I
0
d, I-
z 200
v)
w
AROCY L-10
R T ~
366
n U
T m EPOXIDE
DlEPOXlDE
[r
O
w
VINYL
a
2100
-
POLYESTER
0 0
50
100
150
MONOMER TEMP. ( " C ) at 150 MPA.S
Fig. 5.4 Relationship between fluid monomer temperature and cured T g in families of thermosetting resins.Higher service temperature is normally associated with increasing processing difficulties.
upper right quadrant. AroCy L-10, derived from an asymmetric bisphenol, breaks the pattern and can be used as a 250°C T g resin or as a reactive diluent of 120mPas viscosity (Fig. 5.5).
- 21 0
0
20
40
60
80
1
10
AROCY L-10, WEIGHT %
Fig. 5.5 The asymmetric structureof AroCy L-10 disrupts crystallinity, permitting optional use of this ring-forming resin as a reactive diluent.AroCy numbers are the % cyanate conversion of prepolymers.
but predissolving in 2-6 phr (parts per hundred resin) alkyl phenol, e.g. nonyl or dinonyl phenol, forms stable liquid packages which 5.4.2 CURE CATALYSTS are readily miscible (Shimp, 1988). The alkyl Catalysts are not required to cure BMI resins at phenol provides the active hydrogen co-catatemperatures above 200"C, but several types lyst and can serve as a monofunctional provide effective cure acceleration. Tertiary reactant to increase conversion and resistance amines, imidazoles and free radical generators to boiling water at marginal cure temperatures are noted by Zahir (1978). Boyd (1987) (Fig. 5.6). Extension of this principle to AroCy describes the preferred latency of tri- XU-366 enables this monomer to convert satisphenylphosphine and its phosphonium halide factorily at 121°C for use with high modulus polyethylene fibers (Shimp, 1994a) and with derivatives as prepreg catalysts. Cyanate esters require catalysis to cure at composite tools. practical rates. Copper (most active at low temperatures) and cobalt (latent) acetylaceto5.4.3 TOUGHENING TECHNIQUES nates provide 295% conversion within 2-6 h at post cure temperatures in the range of Concentrated effort over the last decade has 200-250°C. Metal coordination catalysts in produced composite toughening techniques general are difficult to solubilize in neat resins, which satisfy damage tolerance requirements
Matrix and composite properties 105
I
".
0
L
4
2 6 NONYLPHENOL CONC. (phr)
I
0
' I
the CAI performance level of 245 ksi was demonstrated (Boyd, 1993a) by combining in situ epoxy extension of o,o'-diallylbisphenol A with thermoplastic polyimide particles. CE monomers and prepolymers dissolve powdered amorphous thermoplastics (Tps) of the polysulfone, polyethersulfone, polyetherimide, polyphenylene oxide and copolyester families, then subsequently phase separate these thermoplasticsduring cure. Co-continuous morphologies are developed at Tp concentrations 215% which increase GI, values in a nonlinear response to concentration (Shimp, 1994a). Lee (1991) describes the development of a CE matrix formulation toughened with polyoxazolidinones, polyethersulfone and copolyester Tp resins. Table 5.3 lists a number of Tp resins used to toughen both CE and BMI resins as well as reactive rubbers used to eliminate microcracking in orbital service.
I
10
20 30 Meq OHIOCN
I
5.4.4 EPOXY RESIN MODIFIERS
40
Fig. 5.6 Increasing concentrations of alkyl phenols in cyanate ester homopolymers increase conversion (numbersat right) for a given cure temperature and increase resistance of 3 mm thick castings to hydrolysis in boiling water. Cure temperature: a: 250°C; A,:210°C; 0 :177°C. (for AroCy B)
of primary aircraft structures and microcrack resistance in earth orbit and cryogenic service. BMI resins earlier required the development of allyl, propenyl and amine functional reactants to achieve >2% tensile elongation-at-break, a minimum requirement for efficient secondary toughening with thermoplastic polyimides. Boyd (1990) describes the development of improved chain extending reactants by coupling 2- or 4-propenyl phenol with diepoxides. Incorporation of thermoplastic polyimide fine particles of 5-15 pm diameter, described by Boyd (1991a),increased the compression-afterimpact (CAI) performance of BMI composites to 2276 MPa (240 ksi). Further toughening to
Epoxy resins derived from epichlorohydrin co-react with CE resins at equivalent ratios of up to 1.2 epoxides per monomer cyanate. Hybrids with typical epoxy weight fractions of 50-70'/0 develop Tg values in the 180-200°C range and retain dielectric constants 13.1 with loss tangents generally below 0.010 (Shimp, 1992). The use of diepoxides to toughen BMI resins via chain extension of alkenyl (bis)phenols was described in the previous section. Epoxides can also react with the secondary amine formed by the Michael addition of aromatic diamine to the BMI maleic double bond. 5.5 MATRIX A N D COMPOSITE PROPERTIES
The following acronyms will be used to denote BMI resin system properties in figures used thoughout this section: BMI-MDA = the reaction product of 4,4'-bismaleimidodiphenyl methane (molar excess) with methylene dianiline. BMI-DAB = the equimolar reaction product of
106 Speciality matrix resins Table 5.3 Thermoplasticand reactive rubber tougheners
Classification
Product
Used with
Supplier
ICI/Mitsui Amoco General Electric General Electric Ciba
Soluble T,"
Polyethersulfone Polysulfone Polyetherimide Polyphenyleneoxide Polyimide
Victrex 5003P Udel P-1700 Ultem lOOO(P)
203 175 215
PPO
Matrimid 5218
202 300
CE CE CE CE CE
Vitel PE-307
14
CE
Bostik
ATX-013 Hycar ETBN CRS (exp.)
<25 <25 <25
CE CE CE
Echo Resius B.F. Goodrich Dow Chemical
Experimental PAP Series
<25 <25
CE CE
Proprietary National Starch
P-84 Matrimid 5218 1002 D NAT
290 300 85
BMI BMI CE
Lenzing AG Ciba Atochem Corp.
Elastomeric T,
Copolyester Reactive rubbers
Solublea(OH) Soluble"(Epoxy) Preformed Core/Shellb Polysiloxanes
Epoxy functional Maleimide functional Particulate T,"
Polyimide Polyimide Polyamide a
Initially soluble but phase separate during cure. Small particles swell but do not completely dissolve with cure.
Note: Most of this toughening technology is described in patents.
4,4'-bismaleimido diphenylmethane with o,o'-diallylbisphenol A (Matrimid 5292). 5.5.1 MECHANICAL PROPERTIES Table 5.4 compares properties of representative CE and BMI castings. Significant differences are the higher room temperature modulus values of BMI matrices (superior stress transfer to fiber) and higher CE elongation-at-break values.
Although CE resins demonstrate higher temperature onsets of rapid thermal degradation in TGA tests, available long term isothermal aging tests in air indicate superior BMI performance. Boyd (1993b) classifies 1-year CAI retention life of BMI composites as >177 but <205"C. Stenzenberger (1991) rates 2000 h life of BMI castings as >2OO0Cbut <250"C based on retention of shear and flexure strengths. CE/E-glass laminates are rated at 162-180°C for 25 000 h retention of flexure strength at 50% of the unaged values (Shimp, 1989).
5.5.2 THERMAL PROPERTIES
5.5.3 DIELECTRIC PROPERTIES
T , values average about 20°C higher for BMI matrices (Table 5.5). CTE values below T,are comparable while CE resins retain higher char yields as a result of increased aromaticity.
Figure 5.7 ranks the dielectric constant (D,) and dissipation factor (D,) or loss tangent of thermoset matrix resin castings compared with reference thermoplastics. Effects of
Matrix and composite properties 107 RESIN DIELECTRIC PROPERTIES 25 O C
'I,
DK
Df
DK
HOMOPOLYMER
BMI-DAB
AroCy 6 , L AroCy M ,F
AroCy B AroCy M AroCy F XU - 366 POLYETHYLENE
2
1
F 1
PTFE
1
AIR
0
I
I oe5
20 40 60 80 11 0 RESIN CONTENT, Volume %
Fig. 5.8 Effect of reinforcement and concentration on dielectric constant of AroCy M composites. Test data at 25°C and 1 MHz.
tems. Moisture locates primarily in the CE free volume fraction, resulting in less swelling (Fig. 5. 10) than is caused by association with strong dipoles. Plasticization of moisture-conmoisture absorption, test temperature and fre- ditioned matrix castings, compared as a quency are summarized by Shimp (1994b). function of flexural modulus retention at eleLow D , values of CE homopolymers and vated test temperatures, is minimized by the CE/epoxy hybrids are attributed to the sym- low absorption of AroCy M o-methylated CE metrical arrangement of electronegative resin (Fig. 5.11). Hydrolysis of AroCy M cyaoxygen and nitrogen atoms around a central nurate linkages in 121°C steam requires >600 h electropositive carbon atom in these struc- exposure (Fig. 5.12). tures, resulting in weak dipoles. Dielectric constants of CE composites compared with 5.5.5 PROPERTIES OF UNIDIRECTIONAL fiber type and loading are plotted in Fig. 5.8. COMPOSITES D, and loss tangent values of quartz reinforced BMI, CE and epoxy composites are compared Properties of intermediate modulus carbon at four radar bandwidths in Fig. 5.9 (Speak, fiber reinforced BMI composites (Table 5.6) 1991). and CE composites (Table 5.7) indicate good translation of fiber strength for both classes. Damage tolerance ratings based on CAI 5.5.4 MOISTURE ABSORPTION AND EFFECTS results at 6.7 KJ m-l impact energy fall in the CE homopolymers absorb less moisture than 200-345 MPa (30-50 ksi) class, approaching or BMI and TGMDA/DDS epoxy matrix sys- equaling the damage resistance of thermoplastic Fig. 5.7 Thermosetting and thermoplastic resins are ranked for dielectric constant (D,) and dissipation factor (D,) at 1 MHz frequency.
108 Speciality matrix resins
Table 5.4 Mechanical properties of CE and BMI resins M
AroCy L
Matrimid 5292
100
-
-
-
-
100
-
-
100
-
2 0.13
2 0.13
2 0.13
-
88 12.7 3.2 174 25.2
76 11.0 2.7 159 23.0
87 12.6 3.8 187 27.1
82 11.9 2.3 167 24.2
Flexure
Flexure
Flexure
Tensile
3.17 0.46
2.97 0.43 2.35 0.34 175 1.00
3.24 0.47
4.28 0.62 2.42 0.35
AroCy B
Composition (PBW) AroCy B-30 AroCy M-20 AroCy L-10
Matrimid 5292A Matrimid 5292B Nonylphenol Cobalt acetylacetonate
-
AvoCy
-
-
100 85 -
Property of casting" Tensile strength, MPa ksi Tensile elongation, YO Flexure strength, MPa ksi
Young's modulus 25"C,
GPa
msi 149"C, GPa msi 163"C, GPa
msi 204"C,
GPa
G,,,
J m-2
msi in lb in-2
-
2.55 0.37 140 0.80
-
2.28 0.33
-
-
2.00 0.29 170 0.97
-
190 1.08
-
Step-cure with post cure of 2 h at 250°C for CE; 6 h at 250°C for BMI. Data courtesy of Ciba Specialty Chemicals Corp., Performance Poiymers Group.
a
composites. BMI composites have demonstrated hot-wet performance in 177°C rated aircraft. CE composites may attain that goal with AroCy M resin, but insufficient 177°C hot-wet compression data has been published for commercial materials of this class.
KorexTMaramid/phenolic honeycomb core (DuPont) are recommended for use with CE resins catalyzed with copper or cobalt acetylacetonates to eliminate blistering associated with post cures >190°C. (Shimp, 1993; 1994a). 5.6.2 GALVANIC CORROSION
5.6 DESIGN CONSIDERATIONS 5.6.1 SELECTION OF ARAMID FIBER AND CORE
Aramid fiber and core reinforcements for CE composites should be selected from second generation materials wluch absorb <2% moisture in the workplace. Kevlar@aramid fiber and
Both BMI and CE carbon fiber composites have been reported to undergo resin degradation in accelerated galvanic cell tests producing strongly alkaline conditions at cathodic sites (Boyd, 1991b; Olesen, 1991). Figure 5.13 compares the onsets and rates of alkaline hydrolysis (etching) for CE and BMI matrix castings. (See also Boyd, 1991b).
Suppliers of prepreg and other formulated products
109
Table 5.5 Thermal properties of CE and BMI resins
AroCy B
Composition (PBW) AroCy B-30 AroCy M-20 AroCy L-10 Matrimid 5292A Matrimid 5292B Nonylphenol Cobalt acetylacetonate
AroCy M
AroCy
Matrimid
L
5292
100
-
-
-
100
-
2 0.13
-
-
100
-
-
-
100 85
2 0.13
2 0.13
-
Property of casting" HDT, "C Dry Wet
254 197
252 226
249 183
273 217
T,' "C by DMA by TMA
289 257
267 255
270 259
295 273
64
66
64
63
411 41
406 46
408 43
371 29
CTE by TMA, ppm/"C 40 to 200°C TGA at 10"C/min Onset in air, "C Char in N,, Yo Specific gravity at 25°C
1.201
1.151
1.228
1.232
Step-cure with post cure of 2 h at 250°C for CE: 6 h at 250°C for BMI. Data courtesy of Ciba Specialty Chemicals Corp., Performance Polymers Group.
a
Effective design practices for susceptible composites are use of titanium rather than aluminum rivets, placement of a fiberglass reinforced insulating ply and/or modification of CE resin with 55-70% epoxy resin. 5.6.3 MICROWAVE TRANSPARENT COMPOSITES
Composite design for radomes, antennas and advanced stealth structures should utilize low dielectric loss materials (Speak, 1991; Shimp, 1994b; Stonier, 1991a,b). Figure 5.14 summarizes microwave interactions with a radome wall. Reflection weakens returning signals and overheats emitter sources; refraction distorts signal quality; absorption decreases signal
strength and generates destructive heat, limiting power and range. CE composites curing at 121°C (250°F), e.g. Bryte Technologies' EX-1515, are thermally compatible with high modulus polyethylene reinforcement. Such composites are characterized by D, values as low as 2.6 and D, values as low as 0.004 when measured at 10 GHz. 5.7 SUPPLIERS OF PREPREG AND OTHER FORMULATED PRODUCTS
Table 5.8 lists suppliers of BMI and/or CE prepreg, adhesive, syntactic foam, RTM/filament winding systems and chopped fiber reinforced molding compounds formable by compression, injection or transfer processes.
110 Speciality matrix resins 4.0 IIELECTRIC CONSTANT
I I
1 I
3.5
3.0 XBAND 8-12 GHz
KaBAND 26-40 GHz
UBAND 40-60 GHz
J WBANC 75-100 GHz
0.030
TANGENT 0.020
0.0 10
tI
CE
'
0.000 XBAND
KaBAND
UBAND
W BAND
Fig. 5.9 Comparison of typical quartz reinforced radome composites for dielectric loss properties measured at four radar bandwidths. Redrawn from Speak, S.C., Sitt, H and Fuse, R.H.. 1991. Novel cyanate ester based products for high performance radome applications. Int. S A M P E Symp., 36 pp. 336-347.
i P
I +
10'
1o2 1o3 Hours at 25°C & >95% RH
1o4
Fig. 5.10 Changes in 3 mm thick bar volumes during water immersion for a period of one year indicate swelling rates and limits of thermoset resins. The ratio of volume increase to total volume of water absorbed (numbers on right) indicates the fraction of water associated with dipoles. A: BMI-MDA; X: BMI-DAB; 0: TGMDA-DDS; 0:AroCy B; 0 :XU-366
Suppliers of prepreg and other formulated products Table 5.6 Properties of BMI/IM-7 unidirectional composites
Mechanical st rengtk
Rigidite 5250-4"
0" Tensile, MPa (ksi) 25°C 0" Compression, MPa (ksi) 25°C Dry 105°C Wet 149°C Wet 177°C Dry 177°C Wet
Rigidite 5260b
2618 (380)
2691 (390)
1820 (235)
1746 (253) 1346 (195) 1276 (185)
-
-
1310 (190) 966 (140)
0" Compressive modulus 25"C, GPa (msi)
158 (23)
152 (22)
Open hole compression, MPa (ksi) 25°C Dry 177°C Dry 191°C Wet
420 (61) 351 (51) 303 (44)
352 (51) 269 (39) 221 (32)
Compression after impact At 4.5 kJ m-l, MPa (ksi) At 6.7 kJ m-l, MPa (ksi)
248 (36) 214 (31)
380 (55) 345 (50)
Edge delamination, MPa (ksi) 25°C
241 (35)
358 (52)
~~
a
Data courtesy of Cytec. Post cure 6 h at 227°C; 60% fiber vol. Data courtesy of Cytec. Post cure 6 h at 215°C; 60% fiber vol.
Table 5.7 Properties of CE unidirectional composites
Cytec 5245C Reinforcementhre Carbon fiber Max. cure temp., "C
Fiberite 954-2
Hexcel HX-1562
IM-6 210
IM-7 232
IM-7 177
0" Tensile, MPa (ksi)
2439 (356)
2814 (408)
2610 (378)
0" Compression, MPa (ksi) 25"C, Dry 121"C, Wet 132"C, Wet 149"C, Wet
1690 1350 1310 987
(245) (196) (190) (143)
1573 (228) 1331 (193)
1700 (246)
214 (31)
262 (38)
317 (46)
262 (38)
269 (39)
Mechanical strength
CAI, MPa (ksi) At 6.7 kJ m-I Edge delamination, MPa (ksi)
-
-
1290 (187)
-
-
1140 (165) -
-
111
112 Speciality matrix resins
0
°
1
1 O0
79 74
61
Fig. 5.11 (left) Moisture plasticization of cast matrix systems is inversely related to the percentage of dry room temperature flexural modulus retained at elevated test temperatures. : at 149°C wet; : at 177°C wet.
Fig. 5.12 (below) Hydrolysis of unsubstituted CE (bisphenol A dicyanate) homopolymer begins to reduce mechanical properties after 200 h exposure to 121°C steam autoclave at 15 psig. Ortho-methylation is an effective technique for increasing hydrolytic stability of cured CE resins in aggressive environments. 0: AroCy B; 0:AroCy M.
~
AroCy AroCy BMI/
% WEIGHT GAIN
121
AROCY B
91
200 400 TIME, HOURS
0
600
Table 5.8 Sources of formulated/compounded CE and BMI products
Supplier Bryte Technologies Cytec Hexcel Fiberite, Inc. YLA
Prepreg CE BMI, CE BMI, CE CE CE
Adhesive foam
Syntactic compound
compound
CE BMI, CE BMI
CE BMI, CE
CE BMI
-
-
-
-
-
-
BM1,CE
CE
CE
CE
-
-
RTM
Compression molding
AppIications 113 +1.0 AROCYBIEKXY
w H.5
P3
I
AROCYM
-I
0
10
20
30
40
50
60
DAYS IMMERSION IN 20% NaOH AT 50°C
Fig. 5.13 Cured CE and BMI resins hydrolyze (etch) in strongly alkaline solutions, as indicated by the onset of weight loss. Ortho-methylated CE resin and blends with epoxy resin (50/50 blend shown) increase resistance to alkaline environments generated in galvanic cells. 5.8 APPLICATIONS
Toughened BMI/carbon fiber composites have been specified as the principal composite material for F-22 fighter primary and secondary structures (Fig. 5.15).BMI service temperatures are sufficiently high for cowlings, nacelles and thrust reversers of jet engines. CE composites
t
Y
TRANSMISSION Fig. 5.14 Interactions of microwaves with a radome wall. were used to construct EFA (Eurofighter) prototypes and are used in construction of the Dassault Rafale. Both materials are candidates for HSCT (High speed civil transport) use. Principal applications for CE composites (McConnell, 1992) include radomes for military aircraft, fighter aircraft retrofitted with improved tracking systems, skins over phase array radar, weather tracking aircraft radar and missile nose cones. CE prepreg reinforced with high modulus pitch-based carbon fibers are preferred materials for earth orbit service,
Fig. 5.15 F-22fighter constructed with BMI composites. Photograph courtesy of Lockheed.
114 Speciality matrix resins demonstrating low outgassing, microcrack resistance and resistance to lo9rads of ionizing radiation (Willis, 1991). Applications in space include communication satellites, solar arrays, parabolic antennas, optical benches and precision segmented reflectors. BMI film adhesives are employed in jet engine or high speed aircraft sandwich panels where hot-wet service up to 190°C is required. CE film and paste adhesives are used together with syntactic foams in the construction of radomes. BMI molding compounds reinforced with up to 65 wt.% of chopped reinforcements are used to mold ducts, drive sprockets for heated rolls in copy machines, helicopter gear boxes and missile strongback mounting supports. REFERENCES Bargain, M. et al. 1971. US Patent 3 562 223. Boyd, J.D. and D.A. Shimp. 1987. US Patent 4 644 039. Boyd, J.D. and Hon-Son R. 1990. US Patent 4 923 928. Boyd, J.D. 1991a. US Patent 5 037 689. Boyd, J. et al. 1991b. Galvanic corrosion effects on carbon fiber composites. Int. S A M P E Symp., 36,1217-1231. Boyd, J.D. and L.N. Repecka. 1993a. US Patent 5 189 116. Boyd, J.D. and G.E.C. Chang. 1993b. Bismaleimide composites for advanced high temperature applications. Int. S A M P E Symp., 38,357-369. King, J.J., Chaudhari M. and Zahir. S. 1984. Nat. SAMPE Conf., 29 392. Lee, F.W. and K.S. Baron. 1991. US Patent 5 045 609. McConnell, V.P. 1992. Tough promises from cyanate esters. Adv. Comp., May/June pp. 28-37. Olesen, K. 1991. Degradation of graphite/polymer composites in the presence of a corroding metal. Read at the High Temple Workshop, 11, Reno, Nevada, 4 Feb 1991.
Rottloff, G. et al. 1977. US Patent 4 028 393. Shimp, D.A. 1988. US Patent 4 785 075. Shimp, D.A., S.J. Ising and J.R. Christenson. 1989. Cyanate esters: a new family of high temperature thermosetting resins. SPE/Case Western Conf. on High Temperature Polymers and Their Uses, 1, 127-140. Shimp, D.A. and J.E. Wenhvorth. 1992. Cyanate ester-cured epoxy resin structural composites. Int. SAMPE Symp., 37,293-305. Shimp, D.A and M. Southcott. 1993. Controlling moisture effects during the curing of high T cyanate ester/aramid composites. lnt. S A M P f Symp., 38,370-379. Shimp, D.A. 1994a. Technologically driven applications. In Chemistry and Technology of Cyanate Ester Resins (I. Hamerton Ed.) Chap 10. Blackie, Glasgow, pp. 282-327. Shimp, D. and B. Chin. 1994b. Electrical properties and their significance for applications. In Chemistry and Technology of Cyanate Ester Resins (I. Hamerton Ed.) Chap 8. Blackie, Glasgow, pp. 230-257. Speak, S.C., H. Sitt and R.H. Fuse. 1991. Novel cyanate ester based products for high performance radome applications. Int. S A M P E Symp., 36,336-347. Stenzenberger, H.D. 1990.Chemistry and properties of addition polyimides. In Polyimides (D. Wilson, P.M. Hergenrother and H.D. Stenzenberger, Eds) Chap 4. Blackie, Glasgow. Stenzenberger, H.D. et al. 1991. BMI/bis(allylphenoxy phthalimide)-copolymers: improved thermal oxidative stability. Int. S A M P E Symp., 36 pp. 1232-1243. Stonier, R.A. 1991a. Stealth aircraft and technology from World War I1 to the Gulf, Part I. SAMPE Journal, 27(4), 9-16. Stonier, R.A. 1991b. Stealth aircraft and technology from World War I1 to the Gulf, Part 11. S A M P E Journal, 27(5), 9-18. Willis, P.B. and D.R. Coulter. 1991. Applications of cyanate resins to spacecraft composites. Paper read at 8th Int. Con$ Composite Materials, ECCM/VIII, Honolulu, 15-19 July 1991. Zahir, Sheik A-C. and A. Renner. 1978. US Patent 4 100 140.
THERMOPLASTIC RESINS
4
Lars A.Berglund
6.1 INTRODUCTION
posites offer advantages. They have very low Thermoplastic composites form a fairly new toxicity since they do not contain reactive group of materials. Commercial prepreg tape chemicals (therefore storage life is infinite). such as CF/PEEK (carbon fiber/polyether Because it is possible to remelt and dissolve etherketone) and later CF/PPS (polyphenyle- such thermoplastics, their composites are also nesulfide) was introduced in the early 1980s. easily recycled or combined with other recyHowever, as early as 1966, Menges reported on cled materials in the market for molding improved static strength and fatigue resistance compounds. In the aerospace market, composites based when epoxy was replaced by polyamide 6 as on toughened epoxies dominate. The potena composite matrix (Menges, 1966).In the mid tially cheaper manufacturing of thermoplastic 1970s there was interest in CF/PSU (Po1YSu1- composites has not yet been realized to the fone) due to expectations of better processing methods and improved toughness characteris- extent necessary to motivate large-scale investtics. However, solvent resistance was found to ment in new manufacturing equipment. be a problem. Composites later introduced However, for the next generation of aircraft, based on semi-crystalline thermoplastics, such interest in thermoplastic composites is high. as PEEK and PPS, which have been introduced Higher flying speeds require higher temperatures in the materials than the maximum more recently, have excellent chemical resistemperature available from epoxy-based comtance and are superior to epoxy-based posites in use today. Since the release of gases composites in this respect. Enthusiasm for thermoplastic composites is during processing and inherent brittleness are serious disadvantages of thermoset polyimides, generated for, basically, three different reasons. First, processing can be faster than for thermoplastic composites are of great interest. In the automotive market, thermoplastic thermoset composites since no curing reaction composites are used extensively. Matched-die is required. Thermoplastic composites only compression molding of glass mat thermorequire heating, shaping and cooling. plastics (GMT), primarily based on glass Secondly, the properties are attractive, in particular, high delamination resistance and fiber/polypropylene (GF/PP), is common, because it permits fast processing cycles for damage tolerance, low moisture absorption fairly large components. In the established and the excellent chemical resistance of semifield of injection molded components, matericrystalline polymers. Thirdly, in light of als are used with long fibers (5-10mm) in environmental concerns, thermoplastic commolding pellets. This leads to improved mechanical properties compared with materials based 0; shorter fibeis (Truckenmueller Handbook of Composites.Edited by S.T. Peters. Published and 1991). in 1998 by Chapman &Hall, London. ISBN 0 412 54020 7
116 Thermoplastic resins 6.2 MANUFACTURING METHODS
The principles for thermoplastic composites processing are very different from those for thermoset composites. During the processing of thermosets, the polymer is initially a liquid which then solidifies due to the formation of a three-dimensional molecular network from chemical reactions. Thermoplastics are in the solid state before processing because of their high molecular weight. They are heated above their softening temperature during
processing, for conversion into a high-viscosity melt. Shaping then takes place and the material solidifies on cooling. In Table 6.1, different manufacturing methods used for thermoplastic composites are outlined. Further discussion of thermoplastic composites processing is available in Chapter 24 of this book and in previous reviews (Carlsson 1991; Cogswell, 1992; Kausch, 1993). As with thermoset composites, materials with low fiber volume fractions show ease of
Table 6.1 Manufacturing routes for composites based on thermoplastic resins -
Manufact zi ring ro ii te Open mold processes 1. Autoclave
2. Filament winding 3. Folding
Outline of fabrication and processing methods Unidirectional or woven fibers pre-impregnated by the resin (prepreg) are used. Other forms of prepreg have reinforcing fibers in combination with the resin as fibers or as powder. The prepreg layers are stacked on the mold surface and covered with a flexible bag. Consolidation is obtained by external pressure applied in an autoclave at elevated temperature. Prepreg tape or tape with the resin as fibers or powder are wound onto a mandrel at pre-determined angles. Heat and pressure are applied to the tape in order to continuously weld it onto the underlying material. Preconsolidated sheets are heated. Simple fixtures are then used to shape the sheets into the desired geometry.
Closed mold processes A mixture of molten thermoplastic and short fibers is injected into a colder 4. Injection molding (short fibers, 0.1-10 mm) metal mold at very high pressure. The component is allowed to solidify and is automatically ejected. Semi-finished sheets of glass mat thermoplastics are heated and placed in 5. Compression molding the lower part of the mold in a fast press. The press is quickly closed and (short fibers, 5-50 mm) pressure is applied so that the material can flow to fill the mold. Technology is also available where the hot molding compound reaches the mold from an extruder. The same principle as for short fiber materials. 6. Compression molding Continuous fibers require special clamping fixtures for the sheets and can (continuous fibers) primarily be used for simple geometries. A stack of prepreg is placed inbetween two diaphragms (superplastic 7. Diaphragm forming aluminium or polymer film). The diaphragms are fixed whereas the prepreg can move freely. The material is slowly deformed by external pressure and the mold. Prepreg tape or tape with the resin as fibers or powder is pulled through 8. Pultrusion a heated die to form beams or similar continuous structures with constant cross-section geometry. The material is allowed to cool and solidify. Dry reinforcing fibers are placed in the mold. Monomers and/or low 9. Resin injection molecular weight polymer with low viscosity are injected, the reinforcement is impregnated. Polymerization to a high molecular weight thermoplastic occurs by mixing of reactive components and/or thermal activation.
Material forms processing but low stiffness and strength. On the other hand, materials with high fiber content have high stiffness and strength but require slow processing and are difficult to shape into geometrically complicated structures. For high fiber content materials, the high viscosity of a molten thermoplastic usually requires some kind of prepreg fabrication step before final processing. The prepregs may need to be combined into the consolidated, semi-finished sheets before the final processing step. Regular autoclave processing can be used for thermoplastic composites. For most highperformance thermoplastics, however, temperatures have to be higher than the typical 177°C used for epoxy-based composites. Often, the composite manufacturer must purchase a new autoclave if this is the preferred processing route. Autoclave processing of thermoplastics has been modeled (Lee and Springer, 1987). Consolidation of the prepreg layers is an important issue. At a given temperature, sufficient time must be available for the polymer molecules to diffuse from one prepreg layer into the other and form strong physical entanglements (Howes, Loos and Hinkley, 1989). In addition, the air initially present in the material must be displaced. For thermoplastic composites, filament winding has demonstrated good economic potential (Egerton and Gruber, 1988). The major problem is in the welding of filaments or the tape onto the underlying composite layers. Heat has been applied by means of a gas flame, IR, laser beam or simply from a hot metal surface. Pultrusion of thermoplastic composites offers potential for faster processing than with thermoset composites (Astrom, Larsson and Pipes, 1991), due to the absence of exothermal heat generation from chemical reactions. Profiles may also be produced by roll-forming techniques similar to those used in metalworking. The shape of existing profiles can be changed. The low-cost folding technique (GE Plastics, 1990) has been used commercially by Fokker and TenCate in Holland for quite large components of fairly simple geometry.
117
Compression molding of glass mat thermoplastics (GMT) is a wide-spread process of great sigruficance in the automotive industry (Berglund and Ericson, 1994).Resin injection of polymerizing prepolymer molecules of low viscosity is in principle the same process as for thermosets although the chemical reactions lead to increased molecular weight rather than to cross-linking. Such a process does not provide the advantages of infinite storage life materials with low toxicity. Diaphragm forming is a processing route where the problem of low extensibility of prepreg-based materials is addressed (Mallon, O’Bradaigh and Pipes, 1989). 6.3 MATERIAL FORMS
Thermoplastic composites are usually supplied as semi-finished materials, with the exception of resin injection materials. In Table 6.2, material forms for thermoplastic composites are presented. Prepregs of high fiber volume fractions (V, = 0.6) may be prepared by solvent-, melt-, prepolymer- or powderimpregnation of the reinforcing fibers. Solvent-impregnation is limited to amorphous resins with high solubility. Melt-impregnation is a technique successfully developed by IC1 (Cogswell, Hezzell and Williams, 1981) producing high-quality prepreg. The resulting prepreg is considered too stiff, for some processing situations with little drapability in comparison with CF/EP (epoxy) prepreg. This problem is addressed in prepolymer- and powder-impregnated prepreg. One example is the FIT-technology where small tubes containing reinforcing fibers and polymer powder are used (Thiede-Smet, 1989). In addition, commingled weaves (prepregs) are available. The resin is present in the form of fibers which are melted during processing to form a matrix. Composites produced from commingled material forms may have a fairly inhomogeneous distribution of fibers (Olson, 1990). Film-stacking is a simple method often used for preparation of laboratory samples
118 Thermoplastic resins Table 6.2 Material forms for composites based on thermoplastic resins
Material forms
Outline of preparation procedure
Prepregs 1. Solvent impregnated
Reinforcing fibers are impregnated by a mixture of solvent and thermoplastic. The solvent is removed by evaporation. Reinforcing fibers are impregnated by thin molten films to which pressure 2. Melt impregnated is applied. 3. Prepolymer impregnated Low viscosity prepolymers are used to impregnate the fibers. Polymerization to high molecular weight thermoplastic takes place during processing. Fibers are enclosed by thermoplastic powder, either in small tubes 4. Powder impregnated containing reinforcing fibers and powder or by powder particles adhering to the fibers from partial melting in a fluidized bed. Other material forms 5. Film stacked composites Stacks with alternating layers of dry fibers and polymer film are heated and compressed. 6. Fiber hybridized weaves Roving of commingled reinforcing fibers and the matrix in fiber form. Weaves are produced from the roving. and roving Reinforcing weaves are impregnated by prepolymer liquid and 7. Prepolymer liquid and polymerized to a thermoplastic composite. dry reinforcement 8. Semi-finished glass mats Supplied as sheets prepared in belt press by extrusion and melting of films which impregnate fiber mats. Porous sheets are produced from slurry (low fiber content) of fibers and thermoplastic powder in water, by technology similar to manufacturing of paper. Recycled material can be used. Often prepared by pultrusion of unidirectional fibers and matrix 9. Pellets followed by chopping into pellets. Used in injection molding or in (low fiber content) plasticizing and/or compounding unit combined with compression moulding. Recycled material is easily incorporated.
(Hartness, 1982) although the technique has also been used commercially. In the category of materials with low fiber volume fraction (V,=: 0.2), semi-finished sheets of GMT-materials are available. They usually have random, chopped or continuous fiber mat reinforcements. Unidirectional prepreg may be used in order to selectively provide additional stiffness, strength a n d creep resistance. A n interesting step forward is provided by extrusion compounded GMT (Composite Products Inc, 1994; Hoechst AG, 1994). No semi-finished sheets are used, instead a special extruder is used to produce a hot, soft 'cake' constituted of chopped fibers and the polymer matrix, often PP. The cake is placed in a press and molded. The investment in technology is higher than for conventional GMT molding.
On the other hand, the material cost and energy consumption is reduced, greater freedom in materials selection is obtained and recycling is facilitated. Suppliers of thermoplastic composites are listed in Table 6.3. 6.4 THERMOPLASTIC RESINS
Thermoplastics have either amorphous or semi-crystalline structure (Sperling, 1992). The large, chain-like polymer molecules do not show long-range order in amorphous thermoplastics, which may be viewed as polymer glasses and, in the absence of color pigments, are usually transparent. Thermosets are also amorphous. In contrast, crystalline polymers have regions of molecular order. In meltprocessed crystalline polymers, a spherical
Thermoplastic resins 119 Table 6.3 Suppliers of thermoplastic composites
Supplier
Materials
Ba ycomp Burlington, Ontario, Canada
Unidirectional tapes. Matrices PP, HDPE, PA12 PC, PEI, PBT, PES, PPS, PEEK, ABS, PPO. Fibers: glass, carbon, aramid and stainless steel.
CYTEC, Anaheim CA, USA
Commingled yams. Carbon fiber with PEEK, PEKEKK, PA6,6, TPI (Aurum@).GF/PA6,6.
DuPont de Nemours Bad Homburg, Germany and Newark, DE, USA
Prepreg based on Avimid@K,thermoplastic polyimide, and carbon fiber. Sheets laminated of continuous fiber thermoplastic composites or unidirectional discontinuous fibers. Molding compounds of lower fiber content. Matrices: PA6,6, PEKK, PET and others. Fibers: carbon, glass and aramid.
Electrostatic Technology Branford, C, USA
Prepreg fabrication by deposition of polymers in powder form on tow and fabrics. Wide variety of resins and fibers.
GE Plastics Amsterdam, Netherlands and Pittsfield, MA, USA
GMT-materials based on glass fiber mats and PP, PBT, PC and blends PC/PBT. Unidirectional GF/PP.
Hoechst Frankfurt, Germany
Unidirectional prepreg of GF/PP, GF/PA6, GF/PE, CF/PPS, CF/PA6. Pellets > 12 mm for use in plasticating extruder combined with compression moulding.
Huls Marl, Germany
GF/PA12 fabric prepreg.
ICI/Fiberite Monchengladbach, Germany and Laguna Hills, CA, USA
APC-2 (CF/PEEK) prepreg tape and tow and developmental materials, primarily for high-temperature applications.
Porcher Textile Lyon, France
FIT-weaves (Thiede-Smet, 1989).Matrices PA12, PEI, PEEK. Glass and carbon fibers. Enichem, Milano, Italy reportedly produces GF/PP, PET, PBT with FIT-technology.
Quadrax Corp Portsmouth, RI,USA
Prepreg fabrics and unidirectional tape, consolidated sheets. Matrices PA6,6, PMMA, PEI, PPS and PEEK. Carbon, glass and aramid fibers.
Schappe Techniques Charnoz, France
Spun yarns combining reinforcing and matrix fibers for subsequent weaving. Matrices PP, PA6, PA6,6, PPS, PC, PEI, PEEK. Carbon, glass and aramid fibers.
Symalit AG Lenzburg, Schweiz
Glass mat thermoplastic sheets based on GF/PP.
TenCate Advanced Composites Nijverdal, Netherlands and Fountain Valley, CA, USA
Prepreg fabrics and unidirectional tape, consolidated sheets. Matrices PES, PEI and PA12. Carbon, glass and aramid fibers.
120 Thermoplastic resins morphology, termed spherulitic, is often observed (Bassett, 1981). Crystalline lamellae are present within the spherulites although disordered regions exist between and within the lamellae. This is because the large size of polymer molecules inhibits perfect crystallization. The crystalline thermoplastics are therefore more correctly described as semicrystalline, since the degree of crystallinity never reaches 100%. Semi-crystalline thermoplastics can be viewed as two-phase materials with a crystalline and an amorphous phase. To illustrate the difference in behavior of semi-crystalline and amorphous thermoplastics, polyethylene terephthalate (PET),may be used as an example. PET is a thermoplastic polyester which crystallizes fairly slowly. Therefore, upon rapid cooling from the molten state, crystallization can be suppressed and an amorphous polymer is obtained (similar behaviour is shown by PEEK). Samples of PET with different degrees of crystallinity can be produced by changing the conditions of cooling. The shear modulus G' (obtained from dynamic mechanical thermal analysis, DMTA) is plotted against temperature for such samples in Fig. 6.1. The
I
I
I
50
100
I
amorphous sample shows a dramatic drop in modulus at the Tg (glass transition temperature). The drop in modulus for semi-crystalline samples is less dramatic: the higher the crystallinity, the slower the drop. Above T,, material modulus is maintained by the crystalline phase (although strength usually decreases dramatically). Another effect of reduced degree of crystallinity is reduced chemical resistance. A disadvantage with semi-crystalline polymers is the high processing temperature, see Tables 6.4 and 6.5, compared with the heat deflection temperature (see next section, Table 6.11). The melting temperature, Tm,of the crystalline phase must be exceeded during processing, although the maximum use temperature, as for amorphous polymers, is still below T,. Characteristic temperatures of thermoplastics used in applications where only moderate temperatures are experienced are presented in Table 6.4. These materials are available from many different chemical companies, therefore trade names and suppliers are not listed. In Table 6.5, thermoplastics for applications at higher temperatures are listed. These polymers
I I
150 Temperature ("C)
I
I
200
250
Fig. 6.1 Shear modulus (G') compared with temperature for PET of different degrees of crystallinity.
Thermoplastic resins
121
Table 6.4 Characteristic temperatures for thermoplastic resins with T,< 90°C
Polymer type Polyolefin
Chemical name
___
( "C)
Processing temp. ("C)
Crystalline
-10
165
200-240
Crystalline
55
265
270-320
Crystalline
35
180
220-260
Crystalline
70
265
280-310
Crystalline
20
240
260-290
-
~
Polypropylene
T,,,
T ("C)
Structure
___
(PP) Polyamides
Polyamide 6,6 (pA6,6) Polyamide 12 (PA12)
Polyesters
Polyethylene terephthalate (PET) Polybutylene terephthalate (PBT)
Table 6.5 Characteristic temperatures for thermoplastic resins with T, 290°C
Polymer fYPe Polyester Polyarylene ether or sulfide
Polysulfones
Chemical name ~~
~~
Polycarbonate (PC) Polyphenylene sulfide (PPS) Polyarylene sulfide PEEK PEEKK PEKK PEKEKK Polyketone Polysulphone (PSU) Polyethersulfone (PES)
T,
Trade name and supplier
Structure
Lexan, GE Makrolon, Bayer Ryton, Phillips
Amorphous
150
none
280-330
crystalline
90
280
300-340
PAS-2, Phillips
Amorphous
215
none
330
Victrex PEEK, IC1 Hostatec, Hoechst Declar, DuPont Ultrapek, BASF Victrex HTX, IC1
Crystalline Crystalline Crystalline Crystalline Crystalline
143 165 155 175 205
343 365 340 375 385
380400 390-415 380400 400420 420430
Udel P1700, Amoco
Amorphous
190
none
300-350
Victrex 4100G, IC1 Ultrason E, BASF
Amorphous
220
none
300-320
Torlon C, Amoco Torlon AIX-159 Amoco
Amorphous Amorphous
275 290
none none
350400 350400
Amorphous
217
none
335420
Amorphous Crystalline
250 260
none 390
340-360 400420
Processing temp. ("C)
("Ci
~~
Polyamideimides
Polyamideimide (PAI)
Polyimides
Polyetherimide Ultem, GE (PW Polyimide (TPI) Avimid KIII, Du Pont Polyimide (TPI) Aurum, Mitsui Toatsu
122 Thermoplastic resins of static mechanical properties of the resins is given in Table 6.6. Resins with low Tg,such as PP and PA12, have lower modulus and strength. Their fracture toughness is high and valid data according to linear elastic fracture mechanics are difficult to obtain. Among polymers with T, well above room temperature, the modulus is fairly similar. It is controlled by weak physical forces between the molecules. Viscoelastic effects such as creep and stress relaxation during loading will affect the data. Tensile strength varies more widely than modulus between different resins. As a material property it is unfortunately not very reliable. It is sensitive to loading rate, specimen geometry, specimen preparation and the presence of microscopic flaws on the specimen surface. In addition, uniaxial resin tensile strength is different from resin strength in the composite where the stress state is different. Thermoplastics have higher fracture toughness 6.5 PROPERTIES AND DESIGN than epoxy and other thermosets, although CONSIDERATIONS epoxy fracture toughness can be improved by In contrast to thermoset resins, thermoplastics addition of a thermoplastic (Bucknall and can be dissolved and melted. In general, vis- Gilbert, 1989) or other means. Although not coelastic and plastic effects are more apparent from the table, epoxy modulus is usupronounced in thermoplastics. A presentation ally slightly higher than for thermoplastics.
are more expensive and are often termed ’highperformance’ resins. The higher cost of these materials is due to small material volumes, more expensive monomers and more difficult polymerization procedures. Many resins used in injection molding are so called blends, physical mixtures between two thermoplastics. In the field of commercial composite materials, this technology is primarily used for GMT-materials, where composites based on PC/PBT blends are available (Table 6.3). However, for high-performance resins, blending amorphous with semi-crystalline thermoplastics is an interesting route to improved chemical resistance. Most polymer mixtures form immiscible twophase structures although PEEK and PEI may be mixed to form a miscible blend (Crevecoeur and Groeninckx, 1992).
Table 6.6 Mechanical properties of thermoplastic resins Material -__
PP PA6,6 PET PC Amorph. PA a-2) PPS (Ryton) PAS (PAS-2) PEEK PSU (Udel P1700) PES (Victrex) PA1 (Torlon) PEI (Ultem) TPI (Avimid K-111) TPI (LaRC-TPI) EP(thermoset)
Tensile modulus E (GPu) 1.l-1.6 2.5-3.8 2.74.0 2.3-3.0 3.2 3.5 3.2 3.1-3.8 2.5 2.6 2.84.4 3.0 3.5 3.7 2.8-3.5
Tensile strength 0,( M W 3040 50-80 50-70 60-70 100 80 100 90-100 70 80 90-190 105 100 120 40-120
Fract lire toughness G,? (kJ m-2) -
-
1.6 0.5-0.9 -
4.0 2.5 1.9 3.4
3.3 1.5 1.8 0.1-0.5
Properties and design considerations 123 Many mechanical properties of composites are dominated by the influence of fiber modulus, fiber strength and fiber volume fraction. This is usually true for longitudinal tensile modulus and strength as well as flexural modulus and strength. For thermoplastic composites based on AS-4 carbon fiber, typical tensile data are: longitudinal tensile modulus E,= 130 GPa, longitudinal tensile strength (5, = 1950 MPa. In the present context, we are more interested in properties dominated by the matrix and the fiber/matrix interface. One such property is the transverse tensile strength of unidirectional laminates. When a multidirectional laminate is loaded in in-plane tension, the first major damage mechanism is likely to be matrix cracking in the plies with transverse orientation to the maximum load direction. This reduces laminate stiffness and initiates other damage mechanisms such as delamination. In Table 6.7, transverse strength and modulus are presented for different thermoplastic composites. Fiber volume fractions are high, V ,= 0.5-0.6, For composites based on brittle epoxies, typical transverse strength is 40 MPa. Composites
based on LaRC-TPI, J-2, PAS-2 and K-111 show transverse strengths in the range 3 2 4 1 MPa. Otherwise, typical transverse strengths for thermoplastic composites are in the range 60-90 MPa. Toughened epoxy composites also show fairly high transverse tensile strengths, typically around 75 MPa. The use of transverse tension data in failure criteria will lead to conservative estimates. Data are higher for transverse plies in multidirectional laminates (Berglund, Varna and Yuan, 1991). The modulus data for carbon fiber composites in Table 6.7 appear insensitive to small differences in matrix modulus. Variations in fiber volume fraction and transverse fiber modulus between the materials mask any such effect. The detrimental effect of glass fiber as opposed to carbon fiber is apparent from the GF/PA6,6 and CF/PA6,6 data. GF/PP shows very poor performance, probably due to poor fiber/matrix interfacial adhesion (note the low V, ). Interfacial weakness is also likely to explain the low strength for Kevlar / PEKK. For thermoplastic composites based on AS-4 carbon fiber, Table 6.7 can be used to estimate typical data: transverse
Table 6.7 Transverse tensile properties of thermoplastic composites
Material
Transverse modulus Transverse strengfh ET (GPa) (MPa) ~~~
GF/PP GF/PA6,6 CF/PA6,6 CF/Am. PA CF/PPS CF/PAS CF/PAS CF/PSU K49/PEKK CF/PEKEKK CF/PEEK CF/PAI CF/TPI CF/TPI CF/TPI CF/EP CF/EP
Plytron (V, = 0.35) E-glass/Ultramid G30-500/Ultramid AS-4/J-2 AS-4/Ryton T 650-42/Rade18320 AS-4/PAS-2 AS-4/Udel P1700 Kevlar/PEKK LDFTM G30-500/Ultrapek AS-4/PEEK (ICI) T650-42/PAI-696 AS-4/K-III G30-500/’NewTPI’ AS-4/LaRC-TPI AS-4/3501 (thermoset) HTA/6376C (thermoset)
4 8.6 7.2 9 7.6 8.4 8.3 7 6.2 10.3 8.9 7.6 9 8.3
9 9.9
11 48 72 35 72 61 32 59 21 90 80 67 41 59 33 52 75
124 Thermoplastic resins tensile modulus E, = 8.6 GPa, transverse tensile strength (T, = 75 MPa. High interlaminar toughness is desirable since this suppresses the tendency for delamination crack formation during loading. Interlaminar fracture toughness is determined on double cantilever beam specimens (DCB), usually unidirectional materials are used (Whitney, Browning and Hoogsteden, 1982). In Table 6.8 such data are presented. CF/PEEK shows the highest fracture toughness. All thermoplastic composites show higher toughness than the thermoset composites. There is a difference between crack initiation and crack propagation data (Davies, Benzeggagh and de Charentenay, 1987). The data presented here are crack propagation data; crack initiation data are in general much lower. For tough matrices one may question the applicability of the data to design problems. In DCB experiments, the crack opening displacement (COD) is very high, whereas the COD at small central cracks in stiff laminates is much smaller. Local stress fields and damage mechanisms may therefore be different and affect the measured fracture toughness. At present, delamination fracture toughness from DCB tests are therefore preferentially used to compare material. It has been pointed out that composite data are significantly lower than resin data (Hunston, 1984).In tests
on neat resins, energy is absorbed by yielding and other types of damage when the volume of material is relatively large. In a composite, the presence of fibers tends to limit this material volume. For shear strength, no comparable data for different thermoplastic composites appear to be available in the literature. The interlaminar fracture toughness in mode I1 shear loading, Glrc,is higher for thermoplastic than for comparable thermoset composites (Cantwell and Davies, 1993). This also indicates a higher shear strength for the thermoplastic composites. A typical value for the in-plane shear modulus of thermoplastic composites based on AS-4 carbon fiber is 4.8 GPa which is similar to toughened CF/EP systems but slightly lower than for brittle matrix CF/EP composites. Compressive strength is lower for thermoplastic than for thermoset composites (Table 6.9). Most of the thermoplastic composites are in the range 900-1100MPa whereas typical thermoset composite data are 1700 MPa. Fiber misalignment, shear stiffness and strength have been shown to affect compression strength based on plastic kink band formation (Budiansky, 1983). Compression modulus data in Table 6.9 are similar, in support of similar fiber volume fractions for the materials compared. The composite based on PA6,6 has the lowest strength; PA6,6 also has
Table 6.8 Interlaminar fracture toughness of thermoplastic composites Material
Trade name _
CF/Amorph. PA CF/PPS CF/PEEK CF/PEEK CF/PSU CF/PAI CF/PEI CF/TPI CF/TPI CF/EP CF/EP
_
-~ ~
~
AS-4/ J-2 AS-4/Ryton AS-4/Victrex PEEK IM7/Victrex PEEK AS-4/Udel P1700 T650/Torlon AIX-159 T300/Ultem 1000 AS-4/Avimid K-I11 AS-4/ LaRC-TPI AS-4/3501-6 (thermoset) IM7/8551-7 (thermoset)
Fracture toughness qc(kJ~m-2) ~ 1.3 0.9 2.1 2.5 1.2 1.3 0.9 1.8 0.8 0.2 0.5
Properties and design considerations
125
Table 6.9 Compressive properties of unidirectional thermoplastic composites Material
Trade name
Compression strength (MPa)
__
CFDA6.6 CF/Amorph. PA CF/PPS CF/PAS CF/PEEK CF/PEEK CF /PEKEKK CF/PSU CF/PAI CF/TPI CF/EP CF/EP
G30-500 /Ultramid AS-4/ J-2 AS-4 / Ryton AS-4/ PAS-2 AS4/Victrex PEEK IM7/Victrex PEEK AS-4/Ultrapek AS-4/Udel P1700 C-3000/Torlon C AS-4/Avimid K-111 AS-4/ 3501-6 (thermoset) HTA/6376C (thermoset)
700 1100 940 900 1100 1140 1310 1040 1380 1000 1720 1720
Modulus (MPa)
-~
110 -
130 120 120 -
127 -
110 140 130
the lowest creep modulus and yield stress of superior performance to first generation therthe investigated matrices. It is interesting to moset composites (AS-4/3501-6). This is note that AS-4/PEEK and IM-7/PEEK have because the delaminated area due to the roughly the same strength although the IM-7 impact event is more limited for the thermofiber has higher modulus. The smaller diame- plastic composites. However, toughened ter of the IM-7 fiber appears to have a negative epoxy resin composites combined with tougheffect as expected from Euler-buckling consid- ened interlayers between the plies do in erations. general show as good compression strength Compression strength after impact, a mea- after impact as thermoplastic composites. In sure of laminate and material damage fatigue, delamination resistance is higher for tolerance (Dorey, 1989), is presented in AS-4/PEEK compared with epoxy composites Table 6.10. A quasi-isotropic laminate of given (Gustafsson, 1988). However, in uniaxial tenlay-up and geometry is subjected to impact of sion, brittle CF/EP was found to be superior to a certain energy. Internal damage mechanisms both toughened CF/EP and the thermoplastic such as matrix cracking and delamination composite (Curtis, 1987). Claims have been occur in the laminate. The plate is then sub- made that this observation is due to heating jected to compressive load and the stress and effects in the thermoplastic composite specistrain at failure can be determined. The data mens from testing at high frequency (Moore, show that thermoplastic composites have 1991). Table 6.10 Compression strength after impact of thermoplastic composites Material
CF / PPS CF/PEEK CF / PA1 CF/EP CF/EP
Trade name
AS-4 /Ryton AS-4/Victrex PEEK C-3000/Torlon C AS-4/3501-6 (thermoset) AS-4/8551-7 (thermoset)
Compression strength after impact (impact energy) (28 J) (42 1) (571) o (MPa) o (MPa) o (MPa) 221 179 331 310 290 365 345 317 179 145 131 303 -
126 Thermoplastic resins Table 6.11 Glass and heat deflection temperatures for amorphous thermoplastics
Table 6.12 Glass melting and heat deflection temperatures for semi-crystalline thermoplastics
Material
Material
T, ("C)
PP PA6.6 PET PPS PEEK TPI (Aurum)
-10 55 70 90 143 250
PC Amorph. PA PSU PES PA1 PEI EP (thermoset)
T, ("C)
HDT ("C)
150 160 190 220 290 217 200
132 154 175 203 278 200 180
Increased market need for polymer composites with good performance at elevated temperature has generated interest in thermoplastic composites. Materials with continuous use temperatures above 150°C are of particular interest since they perform better than epoxies. One question is how maximum use temperature relates to Tg. In Table 6.11, Tg and the heat deflection temperature (HDT) for amorphous thermoplastics are presented. HDT is determined by subjecting the material to static load (typically 1.8 MPa) and slowly increasing the temperature. HDT is determined as the temperature at which a critical deflection of the sample is obtained. Table 6.11 shows HDT to be 620°C below the T of amorphous poly6 mers. In comparison with room temperature strength, the strength of the composite is significantly reduced above the HDT. The reason
T,,, ("C) 165 265 265 280 343 388
Structure
PA6,6 PEEK PPS PEI PSU
Crystalline Crystalline Crystalline Amorphous Amorphous Amorphous Amorphous Amorphous Amorphous
PA1
PES Am.PA PC
Hydraulic fluid
0 0 0 A 0 A A -
60 75 41 135 160 238
is that molecular mobility in the polymer is increased dramatically as the temperature approaches T . In Table 8.12, similar data to those in Table 6.11 are presented for semi-crystalline thermoplastics. HDT is usually somewhat higher than Tg. However, for some semicrystalline polymers, HDT is below T (as for amorphous polymers). Creep effects 'will be very strong close to and above Tg.For this reason the maximum temperature for continuous service under significant load is unlikely to exceed a temperature of 20°C below T' for semi-crystalline thermoplastics. The primary advantage of crystallinity is therefore chemical resistance. This is apparent from Table 6.13, where chemical resistance for different thermoplastics is indicated in a qualitative way. Semi-crystalline thermoplastics have much
Table 6.13 Chemical resistance of thermoplastics
Material
HDT ("C)
Chlorinated Ketones kydrocarbons 0 0 0 D D 0 D A D
0 = no effect, A = is absorbed, D = is dissolved
0 0 0 A A 0 A A A
Esters H,O abs. (Yo) 0 0 0 A A 0
A A
8 0.5 0.5 1.2 0.9 2 4 0.3 5 -
Applications better chemical resistance than the amorphous polymers. A notable exception is the high water absorption in PA6,6 caused by hydrophilic groups in its chemical structure. The solvent resistance of a large selection of different thermoplastic composites has been reported (Johnston, Towel1 and Hergenrother, 1991). For most thermoplastics, e.g. PEEK, moisture expansion coefficients of the carbon fiber composite may be taken as 0. Thermal expansion coefficientshave been characterized (Barnes et al., 1990) and are similar to epoxy composites. For AS-4/PEEK, the composite density is 1600 kg m-3. For polymers with high Tg, exposure to elevated temperature may lead to increased density not connected with crystallinity but with the amorphous state. The phenomenon is termed physical aging and leads to a more brittle behavior of the polymer (Kemmish and Hay, 1985). Further work is needed to elucidate the importance of physical aging to composite fracture behavior under practical service conditions. For composites processed at high temperatures, residual stresses will also affect fracture behavior. The magnitude of the residual stresses and, consequently, detrimental effects will increase with increasing cooling rate (Manson and Seferis, 1992). Thermoplastic composites can be joined by the same methods as thermoset composites. Bolted joint performance has been compared for thermoset and thermoplastic composites (Walsh, Vedula and Koczak, 1989) with results in favor of thermoplastic composites. With the semi-crystalline thermoplastics, adhesive bonding requires careful surface preparation (Kinloch and Taig, 1987).This is because of the good chemical resistance and limited solubility of these polymers. However, with careful surface preparation, as good adhesive bonds are obtained as with thermoset composites. The thermoplastic nature of the matrix offers another possibility: fusion bonding. Different methods have been compared (Davies and Cantwell, 1993), including hot gas, IR, laser, ultrasonic, vibration, electrical resistance and
127
induction welding. There are difficulties in controlling the processes and very few methods offer the promise of portable equipment. Various bonding technologies for PEEK composites have been compared (Silverman and Criese, 1989).The study favored a technique for fusion bonding with a polyetherimide film. This approach can also be used in repair of damaged structures, since the temperature needed is below the T, of PEEK. 6.6 APPLICATIONS
Thermoplastic composites can be used in similar applications to thermoset composites. The following examples will demonstrate some of the reasons for choosing a thermoplastic composite material. In Europe, the automotive market for GMT composites is significant. Rapid processing by compression molding, cycle times of typically 30s even for large structures, results in cheaper components. In addition, more functions can be integrated into each component compared with sheet metal structures. Bumper beams dominate the automotive market in the US whereas, in Europe, a wider variety of applications are in commercial production. Typical components are subjected to minor loads or impact and surface appearance is not important. Battery trays, beams supporting the hood, seat supports, oil trays, engine shields and even the complete front end have been produced for Volvo, Volkswagen and others. In the aerospace market, DuPont has supplied thermoplastic polyimide composites to the prototype programs for the F-22 fighter aircraft. In the supersonic civil aircraft program, the same thermoplastic polyimide is considered for wing skins. The main reason for this particular thermoplastic composite is a continuous high maximum temperature. AS-4/PEKK (unidirectional discontinuous fibers) was used by Bell Helicopter Textron in a V-22 tiltrotor thermoplastic wing rib for better open hole compression behavior than thermoset composites for high proportions of
128 Thermoplastic resins Bike helmets are also produced (Fig. 6.2). A semi-finished sheet is heated by IR, transfered to heated dies where a selective clamping system is used to hold the sheet. Forming and consolidation pressure is applied and the mold is cooled before demolding. Hoechst in Germany present applications made by winding, a pressure vessel, tubes, sealings and support rings (Fig. 6.3). Matrices include PA6 and PPS with carbon and glass fibers. In the field of biomedical applications, such as hip prostheses, semi-crystalline thermoplastic matrices offer good potential due to their chemical stability (Williams and Fig. 6.2 Bike helmet, an example of a commercial McNamara, 1987). application of thermoplastic composites (Courtesy: DuPont Europe).
45" ply angles (probably for decreased tendency to delamination). A female steel tool and matched-die press forming was used (Chang, 1992). A stretch forming process was used to fabricate C-section curved fuselage frames for Boeing Helicopter from AS4/PEKK (Chang, 1992). TenCate in Holland supplies materials for Fokker Special Products. Landing flap ribs and impact resistant ice-protection plates are produced for the Dornier 328 aircraft.
REFERENCES Astrom, B.T., P.H. Larsson and R.B. Pipes. 1991. Experimental Investigation of a Thermoplastic Pultrusion Process. 36th Int. SAMPE Symp., pp. 1319-30. Barnes, J.A., Sims I.J., Farrow G. et al. 1990. Thermal Expansion Behaviour of Thermoplastic Composite Materials. J Thermoplastic Composites 3: 66-80. Bassett, D.C. 1981. Principles of Polymer Morphology. London: Cambridge University Press. Berglund, L.A. and Ericson M.L. 1994. Glass Mat Reinforced Polypropylene. In Polypropylene: Structure, Blend and Composites. Vol 3: Composites, ed. J, Karger-Kocsis. London: Chapman and Hall. Berglunh L.A., Varna J. and Yuan J. 1991. Effect of Intralaminar Toughness on Transverse Cracking Strain in Cross-Ply Laminates. Adv. Comp. Mater. 1:225-34. Bucknall, C.B. and A.H. Gilbert. 1989. Toughening of a Tetrafunctional Epoxy Resins Using Polyetherimide. Polymer. 30: 213-18. Budiansky, B. 1983. Micromechanics.Computers and Structures. 16 6-10. Cantwell, W.J. and Davies P. 1993. Short-Term Properties of Carbon Fibre PEEK Composites. In Advanced Thermoplastic Composites, ed. Kausch, H-H. pp. 173-191. Munich: Hanser. Carlsson, L.A., ed. 1991. Thermoplastic Composite Materials. Composite Materials Series, Vol 7. New York: Elsevier. Chang, I.Y. 1992. Thermoplastic Matrix Composites Development Update. 37th Int SAMPE Symp.
I .:'i !
1
Fig. 6.3 Thermoplastic composite applications made by winding: pressure vessel, tubes, sealing and support rings. (Courtesy: Hoechst AG)
References 129 Available from DuPont as report entitled Overview of Thermoplastics Composites Technology. Cogswell, EN. 1992. Thermoplastic Aromatic Polymer Composites. Oxford: Butterworth-Heinemann. Cogswell, EN., Hezzell D.J. and Williams P.J. 1981. US Patent 4 549 920. Composites Products Inc 1994. Brochure of CPI process. Winona, Minnesota, USA. Crevecoeur, G. and Groeninckx G. 1992. MeltSpinning of in situ Composites of a Thermotropic Liquid Crystalline Polyester (TLCP) in a Miscible Matrix of Polyether etherketone (PEEK) and Polyether imide (PEI). Polymer Composites 13: 244-50. Curtis, P.T. 1987. In Investigation of the Tensile Fatigue Behaviour of Improved Carbon Fibre Composite Materials. 6th lCCM and 2nd ECCM Proc., Vol 4. pp. 4.544.64. London: Elsevier Applied Science. Davies, P., Benzeggagh M.L. and de Charentenay F.X. 1987. The Delamination Behavior of Carbon Fiber Reinforced PPS. SAMPE Quarterly 19: 19-24. Davies, I? and Canhvell W.J. 1993. Bonding and Repair of Thermoplastic Composites. In Advanced Thermoplastic Composites, ed. Kausch, H-H, pp. 337-366. Munich: Hanser. Dorey, G. 1989. Damage Tolerance and Damage Assessment in Advanced Composites. In Advanced Composites, ed. 1.K Partridge, pp. 369-98. London: Elsevier Applied Science. Egerton, M, and Gruber M. 1988. Thermoplastic Filament Winding Demonstrating Economics and Properties Via In Situ Consolidation. 33rd lnt. S A M P E Symp., pp. 3546. GE Plastics, 1990. Technopolymer Structures, Design and Processing guide. TPS-400A (6/90)RTB. Gustafsson, C-G. 1988. Initiation and Growth of Fatigue Damage in Graphite/Epoxy and Graphite/PEEK Laminates. PhD thesis 88-9. Royal Inst of Techn, Stockholm, Sweden. ISSN 0280-4646. Hartness, J.T. 1982. Polyether-etherketone Matrix Composites. 14th Natl SAMPE Symp., pp. 2643. Hunston, D.L. 1984. Composite Interlaminar Fracture: Effect of Matrix Fracture Energy. Composites Technology Review 6: 176-80. Hoechst AG. 1994. Compela product brochure. Frankfurt am Main, Germany. Howes, J.C., Loos A.C. and Hinkley J.A. The Effect of Processing on Autohesive Strength Development in Thermoplastic Resins and
Composites. In Advances in Thermoplastic Composite Materials. ed. G.M. Newaz, ASTM STP 1044, pp. 3349. Philadelphia: American Society for Testing and Materials. Johnston, N.J., Towel1 T.W. and Hergenrother P.M. 1991. Physical and Mechanical Properties of High-Performance Thermoplastic Polymers and their Composites. In Thermoplastic Composite Materials. ed. L.A. Carlsson, Composite Materials Series, Vol 7, pp. 27-71. New York: Elsevier. Kausch, H-H, ed. 1993. Advanced Thermoplastic Composites. Munich: Hanser. Kemmish, D.J. and Hay J.N. 1985. The Effect of Physical Ageing on the Properties of Amorphous PEEK. Polymer 26: 905-12. Kinloch, A.J. and Taig C.M. 1987. The Adhesive Bonding of Thermoplastic Composite. J . Adhesion 21: 291-302. Lee, W.I. and Springer G.S. 1987. A Model of the Manufacturing Process of Thermoplastic Matrix Composites. J. Comp Mater. 21: 1017-55. Manson J-A.E. and Seferis J.C. 1992. Process Simulated Laminate (PSL): A Methodology for Internal Stress Characterization in Advanced Composite Materials. J. Comp Mater. 26: 405-31. Mallon, P.J., O'Bradaigh C.M. and Pipes R.B. 1989. Polymeric Diaphragm Forming of Continuous Fibre Reinforced Thermoplastic Matrix Composites. Composites 20: 48-56. Menges, G. 1966. The reinforcement of plastics. Kunststoffe 56: 818-23. (In German). Moore, D.R. 1991. In Thermoplastic Composite Materials. ed. L.A. Carlsson, Composite Materials Series, Vol 7, pp. 331-69. New York: Elsevier. Olson, S.H. 1990. Manufacturing with Commingeled Yarns, Fabrics and Powder Prepreg Thermoplastic Composite Materials. 35th Int SAMPE Symp., pp. 1306-19. Silverman, E.M. and Criese R.A. 1989. Joining Methods of Graphite/PEEK Thermoplastic Composites. SAMPE Journal 25: 34-38. Sperling, L.H. 1992. Introduction to Physical Polymer Science. New York: John Wiley & Sons. Thiede-Smet, M. 1989. Study of Processing Parameters of PEEK/Graphite Composite Fabricated with 'FIT' Prepreg. 33rd Int SAMPE Symp., pp. 2086-2101. Truckenmueller, F, and Fritz H.G. 1991. Injection Moulding of Long Fibre-Reinforced Thermoplastics: A Comparison of Extruded and Pultruded Materials with Direct Addition of
130 Thermoplastic resins Roving Strands. Polymer Eng. Sci. 31: 1316-29. Walsh, R, Vedula M. and Koczak M.J. 1989. Comparative Assessment of Bolted Joints in a Graphite Reinforced Thermoset vs Thermoplastic. S A M P E Quarterly 20: 15-19. Whitney, J.M., Browning C.E. and Hoogsteden W. 1982. A Double Cantilever Beam Test for Characterizing Mode I Delamination of
Composite Materials. J. Reinforced Plas. and
Comp. 1:297-313. Williams, D.F. and McNamara A. 1987. Potential of Polyetheretherketone (PEEK) and CarbonFibre-Reinforced PEEK in Medical Applications. J. Muteu. Sci. Lett. 6 : 188-90.
FIBERGLASS REINFORCEMENT*
7
Dennis J. Vaughan
7.1 INTRODUCTION
(2300°F). The molten glass then flows directly to the fiber-drawing furnace in a direct melt The history of glass is ancient, but its engiflow process (Fig. 7.1) or into a marble making neering scientific development is recent. machine. These marbles can be sorted and can Glass was first produced some 4000 years ago, probably in Egypt in the furnaces used to eventually be remelted and drawn into fibers. Continuous glass fibers are produced when produce pottery. Its first application was as a molten glass from the fiber-drawing furnace is form of adornment in jewelry and as added gravity fed through numerous tiny openings decoration to vases and drinking vessels. in a platinum alloy tank called a bushing The use of glass in fiber form dates back to (Fig. 7.2). The droplets of molten glass that the early seventeenth century when the extrude from each of the bushing’s openings Venetians utilized it to create specialized (Fig. 7.3) are gathered together, mechanically gowns. However, commercial fiberglass did attenuated to the correct dimensions, passed not become a reality until in 1939 the joint through a water spray and over a revolving research efforts of Owen-Illinois and Corning belt that applies a protective and lubricating Glass Works, resulted in the formation of coating known as a size or binder. The fibers Owens-Corning Fiberglas Corporation. are then gathered together in a suitably Textile fiberglass has now grown into a shaped shoe to form a bundle called a strand multi-million dollar industry. Glass fiber can which is wound onto a core at approximately be obtained as a continuous fiber on staple or 190 km/h (120 mile/h). This package of fibers discontinuous fiber. Both forms are made by is then dried or conditioned prior to further the same manufacturing process until the fiber processing and eventually sold as a continudrawing operation. ous filament yarn. Staple fibers are produced by passing a jet of 7.2 FIBERGLASS PRODUCTION air across the openings at the base of the bushing, which pulls individual fibers of The production of glass fibers starts with the approximately2040 cm (8-15 in) long from the dry mixing of silica sand and limestone, boric molten glass that is extruding from each openacid and a number of other products such as ing. These filaments are collected on a rotating clay, coal and fluorspar. These materials are vacuum drum, sprayed with size and gathered melted in a high-refractory furnace, the temperinto a strand. This package of filaments is again ature of the melt being dependent on the glass conditioned or dried prior to processing into a composition, but is generally about 1260°C specific product for further use. Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
* Significant portions of this article appeared in Handbook of Composites, 1982, (G. Lubin ed.) Van Nostrand Reinhold, New York.
132 Fiberglass reinforcement
,
,
RAW MATERIALS LIMESTONE SILICA SAND BORIC ACID ,CLAY ,COAL, F,LMRSPAR
I
U
I
FORYUUTION
I
I
W DECORATIVE AND INDUSTRIAL YARN
r
-1
TWISTING
INSPECTION AND WEIGHING
OVEN HEAT TREATING STRAND CHOPPING
W
7 7 7 -
PACKING
Fig. 7.1 Direct-melt fiberglass manufacturing process.
J
A capsule view of the fiber glass manufacturing process lor yarn. roving and chopped strand takes the fiber from raw rnalerflals batch slage lo 'inished products
Glass composifion 133
Fig. 7.2 Molten glass flows from tiny orifices in platinum bushings (Courtesy of PPG industries).
Each individual fiber is drawn from the bushing opening and must be controlled so that reproducible filaments, strand dimensions and properties are obtained. This control is achieved by the regulation of the melt viscosity, temperature and drawing speed. It is possible, therefore, to obtain a large number of filament diameters by varying the number of openings in the bushing and the drawing conditions. As demand has dictated over the years, the fiberglass industry has established a number of standard filament diameters (Table 7.1). 7.3 GLASS COMPOSITION
Glass is generally defined as an amorphous material, being neither solid or liquid. Chemically glass is made up of elements such as silicon, boron and phosphorus that are converted into glass when combined with oxygen, sulfur, tellurium and selenium. The molecular
Fig. 7.3 Molten glass flows from tiny orifices in platinum bushings (Courtesy of Owens-Corning Fiberglas Industries).
Table 7.1 Fiberglass filament designations
Filament designation B C D DE E G H K
Filament diameter i n ~ l O - ~ 1.5 1.8 2.1 2.5 2.9 3.6 4.2 5.1
Pm 3.8 4.5 5 6 7 9 10 13
arrangement is conducive to formation of an intricate three-dimensional network of oxygen tetrahedra with a silicon atom in the middle, bonded to each oxygen atom. Silicon by itself requires extremely high temperatures for liquefaction. Therefore, other elements are added to the mix to reduce temperatures and to produce
134 Fiberglass reinforcement a viscosity in the molten glass that will allow easy drawing. A number of glass compositions are available depending on the properties desired from the resulting fibers (Table 7.2). A-glass: A high alkali or soda glass is made into fibers for use in applications where good chemical resistance is needed. E-glass: A low alkali glass, based on aluminum borosilicate. This glass possesses excellent electrical insulation properties and is the premium fiber used in the majority of textile fiberglass production. C-glass: A material based on soda borosilicate that produces a fiber that offers excellent chemical resistance. S-2 glass: This glass is made up of magnesium, aluminum silicate and offers higher physical strength. Fibers produced from -~ perthis glass have an approximate cent strength improvement Over those of E-glass composition. D-glass: This fiber made from a low dielectric composition has dielectric loss properties 'Onstant Of 3.8 at mc s-l) superior to that of E-glass (6.O at 1mc s-l). R-glass: A 'Pecial glass that produces fiber that is alkali resistant and is used in reinforcing concrete. Low K: An experimental fiber produced to
0
0
improve dielectric loss properties in electrical applications (similar in performance to D-glass). Hollow fiber: A special glass whose fibers are tube-like or hollow; the material has specific applications in reinforced aircraft parts where weight could be significant. Te glass: A Japanese manufactured S-glass, for higher strength structural application.
7.4 FIBERGLASS PROPERTIES
The composition of the original glass melt probably plays the biggest role in determining the properties of the fiberglass. The continuing widespread use of fiberglass in numerous and diverse applications can be directly related to its inherent unique properties (Table 7.3). The suppliers of these materials in the USA are shown in Table 7.4. 0
High tensile strength: Fiberglass has an exceptionally high tensile strength cornpared with other textile fibers. Its strength to weight ratio exceeds steel wire in applications.
0
Heat and fire resistance: Because fiberglass is inorganic it does not burn or support combustion. Chemical resistance: Fiberglass has excellent resistance to most chemicals and is impervious to fungal, bacterial or insect attack.
0
Table 7.2 Fiberglass compositions (wt.%)
Components
Grade of glass
___
~____._
Silicon oxide Aluminum oxide Ferrous oxide Calcium oxide Magnesium oxide Sodium oxide Potassium oxide Boron oxide Barium oxide Miscellaneous
A (high alkali)
C (chemical)
E (electrical)
72.0 0.6
64.6 4.1
54.3 15.2
-
-
-
10.0 2.5 14.2
13.2 3.3 7.7 1.7 4.7 0.9
17.2 4.7 0.6
0.7
8.0
s (high strength) 64.2 24.8 0.21 0.01 10.27 0.27 0.01 0.2
Fiberglass properties 135 Table 7.3 Properties of fiberglass
Grade of glass -
A
C
E
5
-
Physical properties Specific gravity Mohs hardness Mechanical properties Tensile strength, psi x lo3 (MPa) At 72°F (22°C) At 700°F (371°C) At 1000°F (538°C) Tensile modulus of elasticity at 72°F (22"C),psi X lo6 (GPa) Yield elongation, % Elastic recovery, % Thermal properties Coefficient of thermal linear expansion, OF-' x lo4 ("C-1) Coefficient of thermal conductivity, Btu in h-' f f 2 OF-' (Wm-' K-I 1 Specific heat at 72°F (22°C) Softening point, "F ("C) Electrical properties Dielectric strength, V/mil Dielectric constant at 72°F (22°C) At 69 Hz At lo6 Hz
2.50 -
2.49 6.5
2.54 6.5
2.48 6.5
440 (3033)
440 (3033)
500 (3448) 380 (2620) 250 (1724)
665 (4585) 545 (3758) 350 (2413)
-
-
-
-
10.0 (69.0) 4.8 100
10.5 (72.4) 4.8 100
12.4 (85.5) 5.7 100
4.8 (8.6)
4.0 (7.2)
2.8 (5.0)
3.1 (5.6)
-
-
-
0.176
1340 (727)
0.212 1380 (749)
72 (10.4) 0.197 1545 (841)
-
-
498
-
-
-
-
6.9
7.0
5.9-6.4 6.3
5.0-5.4 5.1
Dissipation (power) factor at 72°F (22°C) At 60 Hz At lo6 Hz
-
-
0.005 0.002
0.003 0.003
Volume resistivity at 72°F (22°C) and 500 V DC, ohm-cm
-
-
1015
10'6
Surface resistivity at 72°F (22°C) and 500 V DC, ohm-cm
-
-
1013
1014
Optical properties Index of refraction
-
-
1.547
1.523
Acoustical properties Velocity of sound, ft/s (m/s)
-
-
17 500 (5330)
19 200 (5850)
0
Moisture resistance: Because fiberglass does not absorb water, it neither swells, stretches nor disintegrates. Fiberglass does not readily rot and continues to maintain its mechanical strength in humid environments.
0
Thermal properties: Due to its low coefficient of thermal linear expansion and high coefficient of thermal conductivity, fiberglass exhibits excellent performance in thermal environments.
136 Fiberglass reinforcement Table 7.4 Summary of available fibers from manufacturers S uppl ier
Type of glass --
0
OCF
E S 5-2 D R A C Low K Te
X
Hollow fiber
X
PPG
Nitto Boeski
Vetrotex
NEG
X
X
X
X
___
X
__
--
X
X
X
X
X
X X
X X
X
Electrical properties: Fiberglass being nonconductive is an ideal choice for electrical insulation, where designers can make use of the high dielectric strength and low dielectric loss properties.
7.5 FIBERGLASS TYPES
7.5.1 FIBERGLASS ROVING
Fiberglass roving is a collection of parallel continuous strands or filaments. Conventional The user may take advantage of one or more rovings are manufactured by winding together of the above properties in manufacturing rein- the number of single strands necessary to proforced composites. By utilizing both the high duce the required yield (the number of yards of strength and the excellent electrical properties roving weighing one pound). Single strand of glass fiber, the aircraft industry has found roving, as the name implies, consist of a single fiberglass an excellent reinforcement for strand of fiberglass filaments. These filaments are drawn from a bushing that has the correct radome applications. The printed circuit board industry has used number of openings, so that the single strand the combination of electrical properties and will have the correct yield. Rovings are generally manufactured from superior dimensional stability of fiberglass to manufacture circuit boards that can be used G or K filaments (see Table 7.1), larger diameunder the various adverse environmental con- ters are also available. Roving yields can vary from 1800 to 225 yd/lb (276 to 222 Tex). The ditions. In numerous applications glass reinforce- use of the word Tex indicates a unit of fiber ment is chosen because it allows composites to fineness that is assessed by the weight in retain maximum properties in high moisture grams of 1000 m of yarn, the lower the numenvironments. E-glass fibers are the material ber, the finer the yarn. of choice because of their excellent water resistance. (E-glass fibers exhibit only a 1-7% 7.5.2 WOVEN ROVING weight loss when exposed to boiling water for 60 min.) This water resistance helps maintain A number of rovings are woven into heavy, the physical and electrical characteristics of coarse weave fabrics for use in materials that the composite over prolonged time exposure require a quick build-up of composite thickness over relatively large areas, such as various to aqueous atmospheres’. marine products and different types of tooling. Woven rovings are manufactured in weights ranging from 400-1400 g/m2 (1240
Fiberglass types 137 oz/yd2) and thicknesses of from 0.7-1.5 mm wet layup techniques or by impregnating and (0.02-0.04 in) (see Table 7.5). applying low pressure (RTM) (see Table 7.5). The fiberglass rovings can be used in conjunction with polyester resins in hand lay-up 7.5.3 FIBERGLASS MAT ~. techniques that typically be used in the The three basic forms of fiberglass mat, are manufacture of boats. Woven roving reinforced chopped strand, continuous and surfacing or laminates can be made by using conventional veil. Table 7.5 Woven roving and polyester laminate data
Woven roving constructions and properties Count, per in (per em)
Weight, oz/yd2 (g / m2)
Thickness, in (mm)
Weave
5x8 (2 X 3.2)
18.0 (610)
0.031 (0.787)
Plain
5x8 (2 X 3.2)
24.0 (814)
0.038 (0.965)
Plain
5x6 (2 X 2.4)
30.0 (1020)
0.049 (1.24)
Plain
5x8 (2 x 3.2)
36.0 (1220)
0.052 (1.32)
Plain
Laminate properties - 24 oz (814 g ) woven roving reinforced Resin type"
FR Epoxy
FR Epoxy
GI' Epoxy
Resin content, wt.%
49.0
45.6
42.5
Thickness, in (mm)
0.299 (7.59)
0.265 (6.73)
0.263 (6.68)
Cureb
RT
P+T
P+T
Flexural strength, psi (MPa) Cond. A' Cond. D2/100d
34 200 (235.8) 38 800 (267.5)
49 800 (343.4) 44 400 (306.1)
73 100 (504.0) 59 600 (410.9)
Flexural modulus, psi (GPa) Cond. A Cond. D2/100
20.6 X lo6 (14.2) 2.12 X lo6 (14.6)
2.41 X lo6 (16.6) 2.31 X lo6 (15.9)
2.92 X 106(20.0) 2.83 X lo6 (19.5)
Compressive strength, psi (MPa) Cond. A Cond. D2/100
29 700 (204.8) 27 300 (188.2)
25 700 (177.2)
25 200 (173.8)
47 300 (326.1) 40 100 (276.5)
Tensile strength, psi (MPa) Cond. A Cond. D2/100
42 000 (289.6) 40 600 (279.9)
48 900 (337.2) 47 400 (326.8)
51 400 (354.4) 50 400 (347.5)
"FR = Fire Retardant; GP = General Purpose "RT = Room Temperature; P + T = Pressure and Elevated Temperature 'Cond. A = as-received condition dCond. D2/100 = immersion for 2 h in water at 100°C.
138 Fiberglass reinforcement Chopped strand mat is a non-woven material which the fiberglass strands are chopped into 2-5 cm (1-2 in) lengths and evenly distributed at random onto a horizontal plane, bound together with some type of chemical size. The mats so produced weigh from 2.6-12g/m2 (0.7-3.0 oz/ft2)and are available in a variety of widths, 60-230 cm (24-90 in) with 9.6 cm (38 in) width being the most typical. Continuous strand mat is made of unchopped continuous strands of fiberglass deposited and interlocked in spiral fashion. This mat is open and flexible, but due to its mechanical interlocking does not require much application of size to achieve adequate handling strength. Surface mat, or veil as it is sometimes called, is a very thin mat of single continuous filaments used frequently as a decorative surface reinforcing layer in hand lay-up or press molding processes in order to minimize surface defects and to prevent ’weep’ in pressurized wound tanks. 7.5.4 TEXTILE FIBERGLASS YARN
A yarn can be described as an assembly of fibers or strands which can be woven into some form of textile material. The continuous individual strand as it emerges from the bushing opening represents the simplest form of textile fiberglass yarn and is designated ’singles’ yarn. However, to ensure the correct and efficient utilization of this yarn in a weaving operation, additional strand integrity has to be added by twisting the yarn slightly (less than one turn per inch). A number of woven fabrics, however, require heavier yarns than can be drawn from the bushing. These heavier yarns can be achieved by combining single strands using a twisting and plying operation. This simply involves twisting one or more single strands together or subsequently plying together two or more of the twisted strands. A yarn can be defined as having an ‘S’ twist if when it is held in a vertical position,
the spirals conform in shape to the central position of the letter ’S’. Alternatively, a yarn is said to have a ’Z’ twist if the spiral conforms in shape to the central position of the letter ‘Z’. Strands that have a simple twist (greater than one turn per inch) will kink, corkscrew and unravel, because their twist is in only one direction. The plying operation will normally eliminate this problem by countering the twist in a twisted ’singles’ yarn with an opposite twist in the plied yarn. For instance, a ’singles’ yarns which have a ’Z’ twist are plied with an ’S’ twist that will result in a balanced yarn. The twisting and plying operations allow the yarn‘s strength, diameter and flexibility to be varied, which in turn allows scope in producing a wide range of suitable woven fabrics. 7.5.5 TEXTURED YARN
’Singles’ or plied yarn can be textured by using a jet of air directed onto the yarn’s surface which results in random but controlled breakage of surface filaments, producing a general increased loft to the fiber surface. The term used to describe such a process is texturizing or bulking and the degree to which it occurs can be controlled by regulating the air pressure and yarn feed rate. The texturing process opens the fiber bundles resulting in some mechanical damage to the surface filaments. The increased surface area allows higher resin absorption during impregnation and produces low glass-to-resin ratios, resulting in more economical laminates. 7.5.6 YARN NOMENCLATURE
Because of the wide variety of fiberglass yarns produced, it is necessary to have a precise system for yarn identification. The standard fiberglass yarn nomenclature is based on both alphabetical and numerical designations. The first alphabetical letter identifies the glass composition (E-glass), the second letter indicates filament type (C = continuous) and
Fiberglass types 139 the third and fourth letters identify filament ECG 150 4/2 3.8s diameter (E 7 micron) [for details see Table 7.11. E = E-glass; C = Continuous filament; G = The first series of numbers in the numerical Filament diameter designation represents 1/1OOth of the basic 150 4/2 = Four basic strands of 150 1/0 are strand yield, the second number series specitwisted together to form 150 4/04 fies the number of single strands twisted 150 4/2 3.8s = Plying two strands of EGG together and the number of twisted yarns 150 4/04 (using 'S' twist to create balance) plied together. The total number of basic The above yarn contains 8 (4 x 2) basic 150 strands in a plied yarn is found by multiplying strands with a glass yield of 1875 (15 000 + 8) these two figures. The yield of the yarns is yards/pound. obtained by dividing the basic strand yield by The nomenclature and typical properties of the total number of strands in the yarn. A third commercially available weaving yarns can be number combined with either the letter 'S' or seen in Table 7.6. 'Z' (designating the type of twist) will sometimes also be included. For example:
Table 7.6 Commercially available fiberglass weaving yarns
Yarn designation
Yield
Glass system
-~
E-glass Yarns ECD 1800 1/0 ECD 1800 1/2
Tex system
ydhb ~~
Minim um breaking strength ___
-
~.
Tex
Ib
N
2.75 5.5
0.25 0.5
1.11 2.22
-
EC5 2.75 1 x 0 EC5 2.75 1 X 2
180 000 90 000
ECD 900 1/0 ECD 900 1/2
EC5 5.5 1 x 0 EC5 5.5 1 x 2
90 000 45 000
5.5 11
0.5 1.1
1.11 4.89
ECD 450 1/0 ECD 450 1/2 ECD 450 1/3 ECD 450 2/2 ECD 450 3/2
EC5 11 1 x 0 EC5 11 1 X2 EC5111x3 EC5112x2 EC5 11 3 X 2
45 000 22 500 15 000 11 250 7500
11 22 33 44 66
1.1 2.2 3.3 4.4 6.6
4.89 9.79 14.7 19.6 29.4
ECD 225 1/0 ECD 225 1/2 ECD 225 1 / 3 ECD 225 2/2 ECD 225 2/3
EC5 22 1 x 0 EC5 22 1 x 2 EC5 22 1 X 3 EC5 22 2 x 2 EC5 22 2 X 3
22 500 11250 7500 5625 3750
22 44 66 88 132
2.2 4.4 6.6 8.8 14.4
9.78 19.6 29.4 39.1 64.1
ECDE 150 1/0 ECDE 150 1/2
EC6 33 1 X 0 EC6 33 1X 2
15 000 7500
33 66
3.5 7.0
15.6 31.1
ECG 150 1/0 ECG 150 1/2 ECG 150 1/3 ECG 150 2/2 ECG 150 2/3 ECG 150 2/4 ECG 150 3/3 ECG 150 4/4
EC9 33 1x 0 EC933 1 X2 EC9 33 1 x 3 EC9 33 2 X 2 EC9 33 2 x 3 EC9 33 2 x 4 EC9 33 3 x 3 EC9 33 4 X 4
15 000 7500 5000 3750 2500 1875 1667 938
33 66 99 132 196 264 297 528
3.5 7.0 9.0 12.0 18.0 24.0 27.0 48.0
15.6 31.1 40.0 53.4 80.1 107 120 214
Continued on next page
140 Fiberglass reinforcement Table 7.6 Continued
Yarn designation Glass system ECDE 75 1 / 0 ECG 75 1/0 ECG 75 1 / 2 ECG 75 1 / 3 ECG 75 2/2 ECG 75 2/3 ECG 75 2/4 ECH 55 1/0 ECDE 37 1/0 ECG 37 1/0 ECG 37 1/2 ECG 37 1 / 3 ECH 25 1/0 ECK 18 1 / 0
Yield
Minimum breaking strength
Tex system 7500 7500 3750 2500 1875 1250 938 5500 3700 3700 1850 1230 2500 1800
66 66 132 198 264 396 528 90 134 134 268 403 198 275
5.7 5.7 11.4 17.1 22.8 34.2 45.6 9.5 11.2 11.2 22.8 34.2 17.0 23.0
25.4 25.4 50.7 76.1 101 152 203 42.3 49.8 49.8 101 152 75.6 102
s c 5 11 1 x 1 s c 9 33 1 x 2 s c 9 33 2 x 2
22 500 7500 3750
22 66 132
-
-
8.2 16.4
36.5 72.9
ET6 33 1 X 0 ET6 66 1 X 0 ET6 134 1 X 0
14 200 7100 3450
35 70 144
EC6 66 1 X 0 EC9 66 1 X 0 EC9 66 1 X 2 EC9 66 1 X 3 EC9 66 2 X 2 EC9 66 2 X 3 EC9 66 2 X 4 EClO 90 1 X 0 EC6 134 1 X 0 EC9 134 1X 0 EC9 134 1 X 2 EC9 134 1 X 3 EClO 198 1 X 0 EC13 275 1 X 0
S-glass Yarns SCD 450 1/2 SCG 150 1/2 SCG 150 2/2 Textured Yarns ETDE 150 1/0 ETDE 75 1/0 ETDE 37 1/0
7.5.7 FIBERGLASS FABRIC
The effect that fiberglass fabrics have on the properties of composite material is largely dependent on the fabric’s construction, which involves fabric count, warp and filling yarn construction, weight and weave pattern. The fabric’s count is made up of the number of warp yarns (’ends’) per inch (cm) in the width and the number of filling yarns (’picks’) per inch (cm) in the lengthwise direction. The warp yarn is the yarn that lies in the lengthwise (machme) direction of the fabric. The filling yarn is the yarn lying in the crosswise direction of the fabric. The physical parameters of the fabric, such as weight, thickness and tensile strength are directly proportional to the types and numbers of yarns used to weave it (see Table 7.7)
1.24 2.2 4.7
5.52 9.79 29.0
There is a variety of weave patterns (Fig. 7.4) that can be used to combine warp and filling yarns to form a fabric. This weave pattern controls the handling characteristics of the fabric and to a large extent the properties of the composite that uses it as a reinforcement. (a) Plain weave: This fabric is constructed so that one warp yarn passes over and under one filling yarn (and vice versa). While the resulting fabric exhibits the greatest degree of stability in respect to yarn slippage and fabric distortion, this stability is also a function of the fabric count and yarn content. (b) Basket weave: A basket weave fabric has two or more warp yarns interlocking over and under two or more filling yarns. While a basket weave is less stable than a plain weave, it is more pliable and will conform or drape more readily to simple contours.
Fiberglass types 141 Table 7.7 Glass-fabricconstruction and properties: physical properties for loomstate fabrics (without finish) Style
104 106
Count, /cm /in
23.6 x 60 x 22.0 x 56 x
20.5 52 22.0 56
236x138 60x35
Warp yarn, Tex system US system
Filling yarn, Tex system US system
Weave
5 5.5 1x 0
5 2.75 1 x 0
Plain
D 900 1/0 D 1800 1/0 D 5.5 1x 0 5 5.5 1x 0 D 900 1 / 0 D 900 1 / 0 D5.51~2 55.51~0 D9001/2 D9001/0 D5.51~2 55.51~2 D9001/2 D9001/2 5111x2 5111x2
Weight,
g/m2 ozf yd?
Plain
Plain
Thickness, Bveaking mm stvengtlz, in N/5 cm lb/in
19.7 0.58 24.7 0.73
0.030 0.0012 0.038 n.nni - _ -,5-
35.6
350 x 130 40x15 394 x 350 4.5 . x 40 .-
D4501/2
D4501/2
60x57
5111x2 DW1/2 D111x2 D 450 1/2
120
23.6 x 22.8
5 111x 2
5531x2 D9001/2 511 1 x 2 D 450 1/2 5 1 11x 2
Crowfoot
107
0.043 613x 175 0.0017 Mx20 0.051 613x350 0.0020 70x40 0.076 718x700 0.003 82x80 0.076 1077 x 525 0.003 123x60 0.102 1095x1050 0.004 125x120 0.102 1095 x 1050
220
23.6 x 22.8 60 x 58
7 22 1x 0
7 22 1x 0
Crowfoot
E 225 1 / 0
E 225 1 / 0
109.2 3.22
0.0889 0.0035
1096 x 1051 125 x 120
118~19.3 30x49.
5111x2
5223x2 D2253/2
Crowfoot
298 8.78
0.229 0.009
438x5250
D4501/2
23.6 x 13.8 60 x 35 23.6 x 18.5 60 x 47
5111x0 D 450 1/0 5111x0
55.51~0 D 900 1/0 5111x0
Plain
35.6 1.05 48.8 1.44
0.043 0.0017 0.051 0 002
613 x 175 70x20 613 x 350 70 x 40
1Q7
108
1 1 116
341 1070
1080
23.6~18.5 60x47 151Fx95.4 6x39 23.6x25.2 60x64
I
236~22.8
1.05 Plain Plain Plain
Plain
Plain
4.8.8 1.44 71.9 2.12 83.7 2.47 107 3.16
1
50x600
700x613 80x70 0.089 788x1 0.0035 9 D x 1 0.109 1310 x 1180
0.060
0.0024
1 1165
23.6 x 20.5 60 x 52 Continued on next page
5111x2
D 450 1/2
9 331 x 0 G 150 1/0
Plain
124 3.66
0.0043
150 x 135
142 Fiberglass reinforcement Table 7.7 Continued
Count, /cm /in
Warp yarn, Tex system US system
Filling yarn, Tex system US system
7511x0 E 110 l / O 751 1/0 E 110 1/0 6331x0
Plain
DE 150 1/0
7511x0 E 110 1/0 751 1 x 0 E 110 1/0 6331x0 DE 150 1/0
1522
1 9 . 3 ~16.5 49x42 18.2 x 12.7 46x45 23.6x19.7 60x50 9.1x 8.7
1523
28 x 20 -_ -
9331 x 2 CG 150 1/2 9333x2 G 150 3 / 2
Plain
24 x 22 11.0 x 7.9
9331x2 G 150 1 / 2 9333x2 G 150 ...3 .,/ 2
13.4 x 12.6 34 x 32
9331x2
9331x2 G 150 1/2
Plain
G 150 1/2
9333x3 G 150 3/ 3 9331x2 G 150 1/2 7221x2 D 225 1/2 7221x0 E 225 1/0 9334x2 G 150 4/2 9331x2 G 150 1 / 2 9331x3 G 150 1/3 9332x2 G 150 2/2 9684x2 G 75 2/2 9334x4 G 150 4/4 968 1x 0 G 75 1 / 0 9331x0 G 150 1/0 6681x0 DE 75 1/0 9331x0 G 150 1/0 6331x0 DE 150 1/0
Plain
Style
Weave
g/m2 oz/yd2 -. _. -
1560 1501 1504
1526 1527 1528 1543 1557 1564 1581 1582 1583 1584 1588 1614 1652 1665 1674 1675
1677
6.7 x 6.7 17 x 17 17.3 x 12.6 44 x 32 19.3 x 11.8 49 x 30 22.4 x 11.8 57 x 30 9241x2 G 37 1/2 22.4 x 21.3 57 x 54 23.6 x 22.0 60 x 56 21.3 x 18.9 54 x 48 17.3 x 13.8 44 x 36 16.5 x 14.2 42 x 36 11.8 x 5.51 30 x 14 20.5 x 20.5 52 x 52 15.7 x 9.5 40 x 24 15.7 x 12.6 40 x 32 15.7 x 12.6 40 x 32
15.7 x 15.7 40 x 40 Continued on next page
~
9333x3
G 150 3/ 3 9331x2 G 150 1/2 9662x2 G 75 2/2 9331x2 G 150 1 / 2 9241x2 G 37 1/2 9331x2 G 150 1/2 9331x3 G 150 1/3 9332x2 G 150 2/2 9684x2 G 75 2/2 9334x4 G 150 4/4 933 1x 0 G 150 1 / 0 9331x0 G 150 1/0 9331x0 G 150 1/0 9331x0 G 150 1/0 6331x0 DE 150 1 / 0
6331x0
6331x0
DE 150 1/0
DE 150 1/0
Weight,
Rain
Plain
Plain
Plain Crowfoot Crowfoot Plain Satin Satin Satin Satin Satin Len0 Plain Plain Plain Plain
Plain
-
Thickness, Breaking mm strength, in N/5 em lb/in
_ -
.
-
0.132
2285 x 515
0.0052 0.127
o m
261x253 2145 x 2102 245x240
0.127
2 6 8 8 ~2233
167.8 4.95 166.1 4.90 147.8
4.36
0.01)50
125 3.70 403 11.9
0.140 0.0055 0.356 0.014
1400 x 160 x 4595 x 525 x
1970 x 1705 225x195
307x255 1180 135 3500 400
185
0.165
5.45
0.0065
437 12.9 203 6.00 298 8.34 184 5.42 431 12.7 302 8.92 454 13.4 570 16.8 861 25.4 1756 51.8 79.0 2.33 141.7 4.18 141.7 3.50 95.6 2.82 95.6 2.82
0.381 4375 x 4245 0.015 500 x 485 0.178 2185 x 1750 0.007 250 x 200 0.229 5250 x 525 0.009 600x60 0.140 3240 x 525 0.0055 370 x 60 0.406 4375 x 3940 0.016 500 x 450 0.216 2975 x 2885 0.0085 340 x 330 0.356 4595 x 4375 0.014 525 x 500 0.457 5690 x 5165 0.018 650 x 590 0.686 8315 x 7005 0.035 950 x 800 1.27 16 635 x 11820 0.050 1900 x 1350 0.127 657x744 0.005 75 x 85 0.114 2154 x 1926 0.00454 246 x 220 0.114 2154 x 1926 0.0049 180 x 180 0.107 1225 x 832 0.0042 140 x 95 0.107 1225 x 832 0.0042 140 x 95
108 3.20
0.114 0.0045
1225 x 1135 140 x 130
Fiberglass types 143 Table 7.7 Continued . . .
Style
Count, /cni
Filling yarn, Tex system US system
Weave
/in
Warp yarn, 7ex system US system
Weight, g/m’ oz/yd2
1678
15.7 x 15.7
9331x0
9331x0
Plain
108.5
16 x 14
2112
15.7 x 15.4 40x39 23.6 x 22.0 60x56
K 18 1 / 0 7221x0
R 18 1 / 0 7221x0
Plain
71.2
E2251/0
E2251/0
7221x0
5111x0
E225L’O
E4501/0
7221x0 F 225 1/11
c 225 I / i )
_ _ _ ~ _ _ ~ _ ~ ~
-~ ~~~
2il3 2116
23.6 x 22.8 60 x 58
7221x0
2117
26~21.6
7221x0
E2251/0
7221x0 R2251/0
23.20
66x55 23.6~228
5221x0
5221
D2251Jf) 0 2 2 5
2123
60x58 15.7x15.4 80x39 23.6x20.5 60 x 52
D225 l / O
2165 2313 2523 2532
3070 3313
13.6 x 25.2 60 x 64 11.0 x 7.9 28 x 20 6.3 x 5.5 16 x 14
276x76 70 x 70 2.3.6 x 24.4 60x62
5221x0 5221x0 D 225 1/0 7221 x 0 E 225 1 / 0 10 198 1x 0 H 25 1 / 0 10 198 1x 0 H251/0 6 16.5 1 x 0 ~
DE 300 1/0 6 16.5 1 x 0
DE 300 1/0 34 1
371 3733
Thickness Breaking mrn strength, in N/5 cm lb/in
7 1 x7.1
18 x 18 3743 19.3 x 11.8 49 x 30 3783 21.2 x 18.9 54 x 48 4522 9.4 x 8.7 (6522) 24x22 4533 7.1 x 7.1 (6533) 1 8 x 18 Continued on next page
91341x0 G 37 1/0 9 134 1x 0 G 37 1/0 9 134 1x 0 G 37 1/0 9331x2 G 150 1/2 9681x2 G 75 1/2
9.66
Plain Plain
Crowfoot
1839 x 1751
0.014
431 x 41_ o.
0.086
788x700
2.10
0.0034
90x80
80.7 2.38
0.081 0.0032
1225X!i2B 14Ox~
0.102 u.001
1095 x 1050 125 x 120
0.085 0.0037
1182xlfM9 135x115
002 0.004
1095XlW W x 120. rnXlP35
107 1.16
Plain
0.109
20.4 3.18 107
3.16
888
9331x0 G 150 1/0
Plain
9331x0
Plain
225
Plain
80.7 2.38 403 11.9 246 7.25 ~.
0.81 3 0.0032 0.330 0.013 0.254
92.1 2.74 81.4 2.40
0.0788 0.0031
2.62
G 150 1/0 511 1 x 0 D 450 1 / 0 10 198 1x 0 H 25 1/0 10 198 1x 0 H251/0
6 165 1 x 0 DE 300 1/0 6 16.5 1x 0 DE 300 11’0 91341
3.70
Plain Plain
Plain
Plain
crowfoot
G 37 1 91341x0 G 37 1/0 5 22 1x 0 D 225 1/0 9 134 1x 0 G 37 1/0 9331x2 G 150 1/2 9681x2 G 75 1/2
431 12.7
Plain
Crowfoot
8Harness Plain Plain
197 5.80 286 8.45 556 16.4 123.4 3.64 200.7 5.92
0.094
0.0051 0.114 0.0045
n.nin 0.0076 0.0030
9ox13Q 1095x1220 125x140 1226 x 525 140 x 60 5075 x 3370 580 x 385 2625 x 2450 3nn x 2x13
1226xk26 lLPoxlg0 1050x2094 120x125
0.390 5900x3700 0.0134 571x423 0.203 0.008 0.203 0.008 0.406 0.0160 0.130 0.0051 0.188 0.0071
2185 x 1750 250 x 200 5250 x 525 600 x 60 4816 x 4290 550 x 490 1226 x 1130 140 x 129 2382 x 2601 272 x 297
144 Fiberglass reinforcement Table 7.7 Continued ~~
Style
Count, /cm /in -
4700
6060
6581 6781 7500 7520 7532 7533 7544
7626 7628 7629
76281
-
5.5 x 5.1 14x 13 23.6x23.6 60x60 22.4x21.3 57x54 22.4 x 21.3 57 x 54 6.3 x5.5 16 x 14 7.1 x 7.1 18 x 18 6.3 x 5.5 16 x 14 7.1 x 7.1 18 x 18 11.0 x 5.5
~
Warp yarn, Tex system US system
Filling yarn,
9681x0
9681x0
G 37 1/0
G 37 1/0
Weave
Tex system US system -
-
6 8 . 2 7 1 ~ 0 68271x0 l b 0 DE 600 5 / 0 9331x2 9331x2 f;1 9 1/2 G 150 112 9681x0 9681x0 S2C6 75 1 / 0 S2CG75 1 / 0 9682x2 9682x2 G 75 2/2 G 75 2/2 9681x3 9681x3 G 75 1/3 G 75 1/3 9681x3 9681x3 G 75 1 / 3 G 75 1 / 3 9681x2 9681x2 G 75 1/2 G 75 1 / 2 9682x2 9682x2
-
-
-
Rain
147.5
Plain
DH
13.4 x 12.6 34 x 32 17.3 x 12.6 44 x 32 17.3 x 13.3 44 x 34
9681x0 G 75 1/0 968 1 x 0 G 75 1/0 9681x0 G 75 1 / 0
9681x0 G 75 1/0 968 1 x 0 G 75 1/0 9681x0 G 75 1 / 0
44 x 20
G 75 1/0
G 37 1/0 (Tex)
59 x 54 17.3 x 12.6 44 x 32
DE 75 1/0 968 1 x 0 G 75 1/0
DE 75 1 / 0 9 6 8 1x 0 G 75 1/0
Weight, g/m2 oz/ydZ
8ws X HS
Plain Plain Plain Plain 2xlBSK.
Plain Plain Plain
4.30 39.0 1.19
Shaded items may be obsolete or not in all vendor current catalogues
0.1%
2189x1445
0.0077 250x165 0.0019 656.7 x 718 0.048 75x82
301.8 8.90 301.X 8.90 327 9.66 294.3 8.68 254 7.50 193 5.70 610
0.228 0.28 0.0090 0.356 0.014 0.304 0.0120 0.254 0.010 0.203 0.008 0.559
3520x2846 400x325 3320 x 28-36 400 x 325 3940 x 3590 450 x 410 2890 x 2890 338 x 330 2930 x 2765 335 x 316 2010 x 1925 230 x 220 6520 x 7265
183 5.40 203 6.00 211.9 6.25
0.168 0.0066 0.178 0.007 0.178 0.007
1970 x 1750 225 x 200 2190 x 1750 250 x 200 2190 x 1883 250 x 215
0.011
250 x 120
6.85
Crowfoot
Thickness, Breaking mm strength, in a N/5 em 1b/in
9.15 203 6.00
0.90
0.010 0.173 0.0068
340 x 330 2190 x 1750 250 x 200
Fiberglass types 145 Notes to Table 7.7
a
All fabrics listed in Table 7.7 which contain D 225 5 22 yam can also be woven with the original E 2257 22 yarn equivalents. Breaking strength of fabrics in SI units: N was determined by testing a specimen of 5 cm width and is for loomstate fabrics. Finishing reduces fabric strength. This thickness is an estimate for dry ply thickness and is an average of several tests. The manufacturer usually reports a range of thicknesses. To estimate the cured ply thickness, divide the dry estimate by the expected fiber volume, i.e. if the dry thickness is 0.1 mm and the fiber volume is 0.5 then the estimated cured ply thickness is 0.1/0.5 or 0.20 mm. See introduction for methods of converting between weight and volume percent. Styles 4522 (6522) and 4533 (6533) are S-2 glass; Styles 6581 and 6781 are S-glass. Data from Reference 3
Fig. 7.4 (a) Plain weave; (b)basket weave; (c) twill weave; (d) crowfoot satin (four harness weave; (e) eightharness satin weave.
(c) Twill weave: This fabric consists of one or more warp yarns over and under two or more filling yarns in a regular pattern. This produces either a straight or a broken diagonal line in the fabric, which produces greater drapeability and stability. (d) Crowfoot satin weave: This weave pattern has one warp yarn interlocking over three and under one filling yarn in an irregular pattern. The fabric that results is extremely pliable and lends itself to conforming to complex contours. (e) Eight-harness satin weave: A fabric constructed with one warp yarn interlocking
over seven and under one filling yam in an irregular pattern. The resulting fabric is very pliable and readily conforms to compound contours. Because this weave pattern allows comparatively high fabric counts, it contributes maximum strength to composites reinforced by it. (f) Unidirectional fabrics: Fabrics produced with heavy warp yarns and light filling yarns, in either crowfoot or long shaft satin weaves. The filling yarns can also be composed of yarns other than glass. These fabrics offer high strength reinforcements in the heavy yarn direction.
146 Fiberglass reinforcement
(g) Non-woven: Unidirectional fabrics that can be produced by sticking the 'warp' and 'filling' yarns together chemically. Although this chemical bonding contributes to the fabric's stability, these fabrics have a firm hand and do not drape over complex contours. 7.5.8 OTHER WOVEN FORMS
Fiberglass yarns can also be woven into tapes, contoured fabrics, fluted core fabrics and three dimensional fabrics. Tapes are usually narrow fabrics [less than 30 cm (12 in) wide], with a secured selvedge to prevent unravelling. Contoured fabrics are woven into a specific geometrical shape. Fluted core fabrics are two parallel layers of fabric tied together by stringers of woven fabrics so that the cross-section is triangular or rectangular. Three dimensional fabrics are really planar or fabrics woven with yarns in three distinct directions within the fabric plane. That is, yams are interwoven in: (i) the machine direction, (ii) +45" from the machine direction, or (iii) -45" from the machine direction. 7.5.9 MILLED FIBERS
Continuous fiberglass strands can be hammer milled into very short fiber lengths [approx. 2-6 mm ( H 6 - X in) long]. The actual fiber lengths are determined by the diameter of the screen openings through which the fibers pass during milling. These milled fibers are usually used as inert fillers for thermoplastic and thermoset resin systems. 7.6 SURFACE TREATMENTS
7.6.1 ROVINGS
In order for fiberglass rovings to be compatible with processing methods and materials, a chemical size is applied during the basic fiber forming operation. The formulation of this size is designed to hold the individual glass fila-
ments together, lubricate the roving for contact with various processing equipment, and to permit the glass filaments to be thoroughly wetted when combined with other materials. Fiberglass roving sizes consist of polyvinyl acetate, polyvinyl alcohol, or PVA/starch as a film former with the addition of such chemicals as chrome complexes, organosilane antistatic agents and lubricants that impart the desired strand characteristics.The film former aids in adhering the filaments together and giving strand integrity that will reduce filament abrasion during fiber drawing, strand conversion and the end use of the rovings by the fabricator. An organosilane or coupling agent is added, that can react with hydroxyl groups on the glass fiber surface and also possesses one or more reactive groups to react with other materials, specifically those present in thermoplastic and thermosetting resins. The silane produces a form of chemical bridge between the glass surface and the resin matrix. The inclusion of antistatic agents and lubricants improves the softness and choppability of the roving. The selection of a roving size is based on its resin compatibility, processing and the expected end use performance. The largest percentage of rovings are used in processes that require relatively short fiber lengths (1-5 cm, 0.5-2.0 in). This type is used in processes such as spray-up, preform molding, bulk molding compounds and continuous lamination. Rovings suitable for chopping are available in varying degrees of softness; the softer rovings are generally more difficult to chop, but they are recommended for use in applications that require intricate shapes, sharp comers and difficult radii. Conversely, the harder types of rovings are generally recommended for use where the choppability is of prime importance and they are used in simple parts with minimum contours and radii. Single strand rovings that are commonly used in filament winding and pultrusion should have good strand integrity, controlled
Design considerations 147 level of broken filaments, good wettability and uniform processability under controlled tension. S-glass based rovings have been used in applications where the composite needs increased physical properties ( e g filament wound pressure vessels).
The continuous fiberglass strands that will be used in weaving are treated at the bushing with a starch-oil binder; the general formula for such a binder can be a partially or fully dextrinized starch or amylose hydrogenated vegetable oil, a cationic wetting agent, emulsifying agent and water. These sizes or binders are intended to protect the fibers from damage during their formation and subsequent operations of twisting, plying and weaving.
nated Y) and hydrolyzable groups (designated X) in materials with a generic structure X,SiRY. The hydrolyzable groups are intermediaries in the formation of silanol groups that bond to the glass surface, whereas the organofunctional groups are designed for reactivity or compatibility with the polymer to be used by the composite manufacturer. (Table 7.8) The chemical functionability of the coupling agent can determine the resistance to varied environments, chemical, physical and thermal (Table 7.9). There are many functions that may be attributed to coupling agents at the glass-resin interface. They may provide lubrication to protect against abrasion during fabrication. The coupling agent mechanism can protect against stress corrosion as a result of water incursion and can help improve the mechanical and electrical properties of reinforced composites.
7.6.3 FIBERGLASS FABRICS
7.7 DESIGN CONSIDERATIONS
Fabrics directly from the loom still have the original binders that were applied at the yarn bushing; the warp yarns have also been treated with a polyvinyl alcohol solution to help protect them during weaving. These binders and sizes are not compatible with the resins used by the composite manufacturers and must be removed prior to impregnation with other polymers. This is usually accomplished by treating the fabric to carefully controlled time-temperature cycles, which results in complete removal of all organic material. However, after this heat treatment no suitable interface exists between the fiberglass surface and the resin matrices that will eventually be used. An improved interface with resulting good adhesion even in adverse environment conditions can be achieved by the application of a coupling agent to the glass fiber surface. The most commonly used coupling agents in current use are those termed silanes. The coupling agent mechanism of these organofunctional silanes depends on a link between the organofunctional group (desig-
It is only recently that the engineering and design criteria in the combination of such different materials as thermosetting and thermoplastic polymers with reinforcing of glass were properly understood. There are a number of basic considerations that are common to the use of fiberglass as a reinforcing media in the production of composites. The type form, weave pattern, weight, permeability and alignment of the fibers have direct bearing on the mechanical properties of the composite. Also in some instances the chemical, thermal and electrical properties may also be influenced by the judicious choice of reinforcing material (Table 7.10).
7.6.2 FIBERGLASS TEXTILE YARNS
7.7.1 GLASS COMPOSITION
The chemical composition of the glass can have a direct bearing on the fiber properties which can indirectly affect composite performance. It has been noted that 'E' glass composition fibers can contribute 210 000-225 000 psi to the
148 Fiberglass reinforcement Table 7.8 Commercial coupling agents
Organofunctionalgroup
Clzemical structure
Vinyl Chloropropyl
CH,= CHSi(OCH,), ClCH,CH,CH,Si(OCH,),
EPOXY
CH2CHCH20CH2CH2CH2Si(OCH3)3
/”\
CH2 Methacrylate Primary amine Diamine Mercapto Cationic styryl
I
(CH2=C-COOCH2CH2CH2Si(OCH3)3 H,NCH,CH,CH,Si(OC,H,), H,NCH,CH,NHCH,CH,CH,Si(OCH,)3 HSCH,CH,CH,Si(OCH,),
CH,=CHC,H,CH,NHCH,CH~NH(CH~),Si(OCH,),HCl CH3
Cationic methacrylate
c1
I
I
+
CH2=CXOOCH2CH2- N(Me2)CH2CH2CH2Si(OCH& CH3
I
CH,=C
I
Chrome complex
CH2 Titanate Cross-linker Mixed silanes Formulated
I
(CH2=C-C00)3 TiOCH(CH& (CH,O), SiCH,CH,Si (OCH,), C,H,Si(OCH,), + F Melamine resin + C
tensile strength of a composite which they reinforce. Other glass compositions/ ’” and ’2’ glass, have produced even greater strength improvements in reinforced composites.
7.7.2 FILAMENT DIAMETER
A number of different ’E’ glass filament diameters have been evaluated comparing the influence of filament diameter on resulting laminate properties. The fabrics evaluated were woven with the same count and weave pattern as Style 1581, using both plied and singles yarn containing different filament diameters (Tables 7.11 and 7.12).
Design considerations 149 Table 7.9 Effect of coupling agents on the mechanical strength properties of laminates Finish
Loomstate
Volan
A1 72
A174,26030
Polyester laminates reinforced with style 1581 glass fabric Flexural strength, psi X lo3 (MPa) Cond. A 40.4 (278.6) 32.3 (222.7) 73.2 (504.7) Cond. D2/100 25.3 (174.4) 31.8 (219.3) 59.0 (406.8)
68.8 (474.4) 66.1 (455.8)
75.5 (520.6) 77.3 (533.0)
Compressive strength, psi X lo3 (MPa) Cond. A Cond. D2/100
112
23.8 (164.1) 12.5 (86.2)
40.4 (278.6) 12.1 (83.4)
48.0 (331.0) 38.9 (268.2)
47.1 (324.8) 44.9 (309.6)
66.2 (456.4) 58.8 (405.4)
48.1 (331.6)
20.6 (142.0)
49.9 (344.1) 48.8 (336.5)
50.0 (344.8) 51.2 (353.7)
52.2 (359.9) 50.7 (349.6)
Epoxy laminates reinforced with style 1581 glass fabric Flexural strength, psi X lo" (MPa) Cond. A 78.2 (539.2) 76.0 (524.0) 80.9 (557.8) Cond. D2/100 76.7 (528.8) 72.4 (499.2) 83.1 (573.0)
71.3 (491.6) 68.9 (475.1)
85.7 (590.9) 82.4 (568.1)
Compressive strength, psi X lo3 (MPa) Cond. A Cond. D2/100
64.3 (443.3) 58.5 (403.3)
63.4 (437.1) 54.1 (373.0)
70.7 (487.5) 64.7 (446.1)
63.0 (434.4) 62.9 (433.7)
67.5 (465.4) 59.5 (410.3)
Tensile strength, psi x lo3 (MPa) Cond. A Cond. D2/100
56.9 (392.3) 55.8 (384.7)
48.8 (336.5) 47.2 (325.4)
57.8 (398.5) 56.1 (386.8)
50.9 (351.0) 49.4 (340.6)
58.2 (401.3) 52.2 (359.9)
Tensile strength, psi X lo3 (MPa) Cond. A Cond. D2/100
Filament diameter does not appear to play an important role in the determination of the mechanical properties of such plied yarn fabrics as Style 1581 woven with ECG 150 1/2 (EC 9 33 1 x 2 yarn and style) and Style 1281 woven with ECB 150 1/2 (EC 3.8 33 1 x 2 yarn). However, in comparison with commercially available singles yarn Style 7781 woven with ECDE 75 1 / 0 (EC 6.66 1 x 0) yarn, singles yarn Style B 7781 woven with finer filaments [ECB 75 1/0 (EC3 8.66 1 x 0 yarn)] gives laminates with higher flexural strengths, flexural modulus and tensile strengths. Conversely a singles yarn G 7681 woven with coarser filaments [ECG 75 1/0 (EC9 66 1 x 0 yarn)] gives composites with improved compressive strengths.
The effect of variation in the available surface area for stress transfer is affected by the coupling agent-resin interaction. Larger surface areas with poor coupling agents reduces specific shear loading on the interface and so produces high laminate flexural strengths. But with improved interfacial bonding the surface areas effect disappears because other failure modes appear first. Singles yarns produce higher composite strengths due to the smaller angle between fiber axis and load axis in the lower twist yarns. This produces less tensile loading across the interface; the magnitude of this effect will vary with the resin coupling agent efficiency and the function of the twist of the yarn or the angle of the spiral of the filaments.
150 Fiberglass reinforcement Table 7.10 The effect of fabric construction on laminate mechanical properties2
Fabric construction Fabric no.
E1581
51581
YM31A"
Count/in (cm)
57 X 54 (22.4 x 21.3)
57 x 54 (22.4 x 21.3)
57 x 54 (22.4 x 21.3)
Warp yarn Glass system Tex system
ECG 150 1/2 EC933 1 x 2
SCG 150 1/ 2 SC933 1 x 2
MCG 150 1 / 2 MC9 33 1 x 2
Filling yarn Glass system Tex system
ECG 150 1/2 EC9 33 1 x 2
SCG 150 1/2 s c 9 33 1 x 2
MCG 150 1/2 MC9 33 1 x 2
Weave
Long-shaft satin
Long-shaft satin
Long-shaft satin
Laminate mechanical properties
Fabric no. Flexural strength, psi x lo" (MPa)
E1581
S1581
YM31A"
88.6 (610.9)
87.1 (600.6)
89.8 (619.2)
Flexural modulus of elasticity, psi x loh (GPa) Cond. A
3.78 (26.1)
3.79 (26.1)
5.00 (34.5)
Compressive strength, psi x lo3 (MPa) Cond. A
64.8 (446.8)
56.4 (388.9)
61.0 (420.6)
60.0 (413.7)
61.4 (423.4)
64.5 (444.7)
Cond. A
Tensile strength, psi x lo" (MPa) Cond. A a
Data from Reference 4
7.7.3 WEAVE PATTERN
The amount of yarn and the pattern of the weave often determine the handling characteristics of a fabric. If a fabric is too tightly woven, it will not readily drape and conform to various contours during molding; a tight weave can also adversely affect impregnation of the fabric by the resin. If a fabric has a weave pattern that is very open, this can produce a weaker reinforcement due to insufficient fiber, a tendency to distort and in horizontal treater, difficulty in impregnation. A general assumption can be made in the longer the float (Le. the portion of a warp or filling yarn that will extend unbound over two
or more yarns lying 90 degrees to it), the higher the reinforced composite strength. So by extending the length of the float which reduces the interlocking frequency, the resulting composite's strength will be increased. Changing the weave, for example from a plain weave to a crowfoot weave, will also achieve improved composites strength properties (Table 7.13). 7.7.4 GLASS-TO-RESIN RATIO
In general, approximation of physical properties for fiber glass reinforced composites follows the 'rule of mixtures'. The calculation of loads stresses and strains parallel to the fibers
Design considerations 151 Table 7.11 Composite mechanical properties as a function of glass-filament diameter
Plied-yarn fabrics ~-
Fabric style Count/in (cm) Warp yam Glass system Tex system Filling yarn Glass system Tex system Weave Flexural strength, psi x lo3 (MPa) Cond. A Flexural modulus of elasticity, psi x 106(GPa) Cond. A Compressive strength, psi x lo3 (MPa) Cond. A Tensile strength, psi x lo3 (MPa) Cond. A
181 57 x 54 (22.4 x 21.3)
1581 57 x 54 (22.4 x 21.3)
1281 57 x 54 (22.4 x 21.3)
ECE 225 1/3 EC7 22 1 x 3
ECG 150 1/2 EC9 33 1 x 2
ECB 150 l / 2 EC3.8 33 1 x 2
ECE 225 1 / 3 EC7 22 1x 3 Long-shaft satin
ECG 150 1 / 2 EC9 33 1 x 2 Long-shaft satin
ECB 150 1/2 EC3.8 33 1x 2 Long-shaft satin
93.0 (641.2)
88.6 (610.9)
92.4 (637.1)
3.88 (26.8)
3.78 (26.1)
4.06 (28.0)
58.1 (400.6)
64.8 (446.8)
64.9 (447.5)
60.0 (413.7)
64.8 (446.8)
60.0 (413.7)
~.
Singles-yarn fabrics ~ _ _ _ _
Fabric style Count/in (cm) Warp yarn Glass system Tex system Filling yarn Glass system Tex system Weave Flexural strength, psi x 10’ (MPa) Cond. A Flexural modulus of elasticity, psi x lo6 (GPa) Cond. A Compressive strength, psi x 10’ (MPa) Cond. A Tensile strength, psi x lo1 (MPa) Cond. A
G7681 60 x 54 (23.6 x 21.3)
DE7781
B7781
60 x 54 (23.6 x 21.3)
60 x 54 (23.6 x 21.3)
ECG 75 1/0 EC9661x0
ECDE 75 1/ 0 EC6 66 1x 0
ECB 75 1/0 EC3.8 66 1 x 0
ECG 75 1/0 EC9 66 1 x 0 Long-shaft satin
ECDE 75 1/0 EC6 66 1 x 0 Long-shaft satin
ECB 75 1/0 EC3.8 66 1 x 0 Long-shaft satin
92.6 (638.5)
94.9 (654.3)
101.1 (696.4)
3.78 (26.1)
3.90 (26.9)
4.02 (27.7)
76.7 (528.8)
67.7 (466.8)
69.9 (482.0)
64.2 (442.7)
66.3 (457.1)
69.4 (478.5)
152 Fiberglass reinforcement Table 7.12 Mechanical properties of S-glass fabrics-epoxy composites as a function of plied compared with unplied yarn S7682
S1581
Fabric style ~~~
~~
57681
-~ -
-
Finish
Volan A
Volan A
CS-310
Flexural strength, psi x lo3 (MPa) Cond. A
84.8 (584.7)
120.0 (827.4)
116.0 (799.8)
Flexural modulus of elasticity, psi x 10" (GPa) Cond. A
3.10 (21.4)
4.18 (28.8)
5.21 (35.9)
Compressive strength, psi x lo3(MPa) Cond. A
61.3 (422.7)
55.8 (384.7)
73.9 (509.5)
Tensile strength, psi x lo1 (MPa) Cond. A
70.6 (486.8)
87.5 (603.3)
86.4 (595.7)
Table 7.13 Effect of weave pattern on composite mechanical properties Fabric style
7628"
76281h
16-149'
7781d
Finish
Volan
Volan A
Volan A
Volan A
Resin
Polyester
Polyester
Polyester
Polyester
Plies
18
18
12
12
Resin content, wt.%
37.1
36.7
36.5
37.6
Thickness, in (mm)
0.124 (3.15)
0.121 (3.07)
0.120 (3.05)
0.120 (3.05)
Flexural strength, psi X 10' (MPa) Cond. A
53.8 (371.0)
84.7 (584.0)
63.2 (435.8)
87.0 (600.0)
Flexural modulus of elasticity, psi x 1Oh (GPa) Cond. A
3.89 (26.8)
3.41 (23.5)
3.81 (26.3)
3.24 (22.3)
Compressive strength, psi X lo1 (MPa) Cond. A
25.7 (177.2)
57.0 (393.0)
48.0 (331.0)
64.3 (443.3)
Tensile strength, psi x 103(MPa) Cond. A
45.9 (316.5)
59.2 (408.2)
58.7 (404.7)
60.0 (413.7)
Plain weave, data from Reference 3 Crowfoot satin weave, data from Reference 3
li
A 5-shaft satin weave version of Style 7781 An 8-shaft satin weave
Design considerations 153 and perpendicular to the fibers is quite different. Also, the stress transfer across the material boundaries is greatly affected by the degree of adhesion at the reinforcement-resin interface. A more subtle effect is the possibility that wall effects at the surface may alter the matrix orientation and mechanical properties. It must also be assumed that there will be sufficient resin present to fill all of the spaces between the fibers. Internal voids have been shown to produce considerable stress concentrations resulting in premature mechanical failure of the laminate. Woven fiberglass fabrics by virtue of the weave interlocking reduce the effective glass-to-resin ratio below those of filament wound structures. The optimum glass to resin ratio then becomes dependent on weave
pattern, weave density and yarn diameters. The optimum strengths of laminates are obtained with the lowest practical resin content. With coarse plain weave fabrics using hand lay-up, laminates can be produced with resin contents of 36-38%; with other weaves using less twist in the yarn and optimum yarn spacing, excellent physical properties can be achieved in the 25% resin content range. Plain weave fabrics with higher twist yarn will require higher resin contents than low twist singles fabrics of equivalent weave. In filament winding applications parallel strands of fiberglass are wound around a cylindrical mandrel, the resulting cross-section of round filaments provides a close packing, resulting in low resin contents of the final part.
Table 7.14 Laminate mechanical properties as a function of glass-to-resin ratio (Style 181 glass fabric-epoxy resin composites)
Plies
7
9
11
14
13 ~
~
16
18
~
20
___
Resin content, wt.%
55.0
49.2
44.0
35.6
31.2
28.0
22.5
22.2
Thickness, in mm
0.117 2.97
0.126 3.20
0.127 3.23
0.126 3.20
0.125 3.18
0.131 3.33
0.132 3.35
0.147 3.73
Flexural strength, Cond. A psi x103 MPa
45.6 314.4
58.8 405.4
64.8 446.8
83.0 572.3
81.2 559.9
92.1 635.0
87.4 602.6
91.9 633.7
Flexural modulus of elasticity, Cond. A psi x l o h GPa
2.26 15.6
2.62 18.1
2.92 20.1
3.33 23.0
3.82 26.3
4.04 27.9
4.37 30.1
4.64 32.0
Compressive strength, Cond. A psi x lo3 MPa
45.2 311.7
47.9 330.3
44.7 308.2
53.2 366.8
56.9 392.3
52.4 361.3
54.3 374.4
67.8 467.5
Tensile strength, Cond. A psi x 103 MPa
30.8 212.4
35.7 246.2
40.5 279.2
51.3 353.7
53.7 370.3
61.0 420.6
64.7 446.1
65.1 448.9
Tensile load/ply, lb N
514 2286
494 2197
468 2082
498 2215
480 2135
500 2224
476 2117
480 2135
154 Fiberglass reinforcement For example, resin contents of 25-30% with glass fiber content of 70-75% product final composites with high physical strengths. But, if there is any variation from the parallel alignment of the glass fibers, this will reduce the degree of packing and inevitably the optimum glass-to-resin ratio. The weaving of glass fabrics will of course reduce the effective glass-to-resin ratios because of the weave interlocking. The correlation between glass-to-resin ratios and composite strength is well known, the composite's strength increases as resin contents are reduced (see Table 7.14 and Fig. 7.4). 7.7.5 FIBER DISTRIBUTION
Reinforced composites that have the glass strands aligned parallel to each other have their maximum strength and stiffness in the direction of the alignment. This type of parallel alignment is conducive to certain filament winding and pultrusion operations. When the reinforcement is aligned at right angles to itself (half of the strands are laid at
right angles to the other half), the resulting mechanical strengths at either angle is less than that of the parallel alignment. As the distribution of strands varies between 0" to 90" alignment to a +45" to a 45", this further reduces mechanical strength in the primary direction, but shows an increase in the +45"and 4 5 " directions. Of course the yarn distribution can be varied from 0" and 90" directions as part of the fabric design. A so called %balanced' fabric, that is one with equal yarn distribution in the warp and filling directions, will have comparable composite properties in both directions. When fiberglass yarn is woven with yarn primarily in the warp direction with only the minimum amount of filling yarn (enough to give the fabric stability), the result is a unidirectional fabric. Composites reinforced with this type of fabric have the maximum mechanical strength in the direction of the greatest concentration of yarn5. The differences between the mechanical properties, bidirectional and unidirectional reinforced composites, can be seen in Table 7.15.
Table 7.15 Effect of fabric yarn distribution on laminate mechanical properties
Bidirectional (2582)
Fabric type (style) _ _ ~
Unidirectional (7743)
~
Warp
Filling
Warp
Filling
52
48
90
10
Flexural strength, psi x lo3 (MPa) Cond. A Cond. D2/100
84.2 (586.0) 75.5 (520.5)
78.2 (539.2) 68.5 (472.3)
95.0 (655.0) 87.4 (602.6)
23.3 (160.6) 24.1 (166.2)
Flexural modulus of elasticity, psi x loh (GPa) Cond. A Cond. D2/100
4.02 (27.7) 3.87 (26.7)
3.65 (25.2) 3.41 (23.5)
4.95 (34.1) 4.80 (33.1)
2.47 (17.0) 1.46 (10.1)
Compressive strength, psi X lo" (MPa) Cond. A Cond. D2/100
62.7 (432.3) 51.8 (357.2)
63.4 (437.1) 54.5 (375.8)
64.5 (444.7) 51.1 (352.3)
28.9 (199.3) 23.6 (162.7)
Tensile strength, psi X lo3 (MPa) Cond. A Cond. D2/100
57.9 (399.2) 55.7 (384.0)
54.9 (378.5) 50.2 (346.1)
94.0 (648.1) 91.6 (631.6)
12.1 (83.4) 11.7 (80.7)
Test direction Yarn content,
'lo
References 155 Table 7.16 Mechanical properties of polyester laminate reinforced with chopped strand mat Resin content, wt.'%,
69.8
Thickness, in (mm)
0.238 (6.05)
Flexural strength, psi (MPa) Cond. A Cond. D2/100
26 400 (182.0) 30 400 (209.6)
Flexural modulus of elasticity, psi (GPa) Cond. A Cond. D2/100
REFERENCES 0.99 x lob (6.83) 0.94 x lo6 (6.48)
Compressive strength, psi (MPa) Cond. A 27 600 (190.3) Cond. D2/100 23 200 (160.0) Tensile strength, psi (MPa) Cond. A Cond. D2/100
Fiberglass remforcement aligned in a random manner within the polymer matrix (e.g. chopped strand mat) produces composites with fairly uniform mechanical strengths in all directions. However, this also tends to produce composites that have relatively low physical properties in all directions (Table 7.16).
14 100 (97.2) 13 800 (95.1)
1. Pitt, C.F. and Harvey J. 20th Anniversary Technical Conference, SPI Reinforced Plastics Division, 1965, Section 9-C. 2. Knox, C.E. Non-Metallic Materials (SAMPE) 4, 127 1972. 3. Horton, R.C. and Adams, R.G. 21st Annual Conference, SPI Reinforced Plastics Division, 1966, Section 10-A. 4. Knox, C.E. New Horizons in Materials and Processing (SAMPE) 18, 527 1973. 5. Peterson, G.P. Properties of High Modulus Reinforced Plastics, S.P.E.J., 57, 1961 January.
BORON, HIGH SILICA, QUARTZ AND CERAMIC FIBERS
8
Anthony Marzullo
tinuous. Continuous fiber or whisker is capable of being manufactured to an indefinite Boron, high silica, fused quartz and ceramic length. Discontinuous fiber can be chopped fibers are used in demanding industrial, autocontinuous fiber, or a short fiber called staple. motive, electronic, aircraft and aerospace Discontinuous whisker is manufactured to a environments. Boards made of ceramic fibers definite length in batch type processing, while such as alumina, alumina-silica or zirconia are continuous whisker is produced by melt used as supplemental linings to insulate high processes such as the laser-heated floatingtemperature furnaces. Ceramic tiles containzone process. ing alumina and silica fibers protect the Summarizing, there is continuous fiber aluminum skin of USA's space shuttle orbiter (indefinite length), continuous whisker (single during re-entry, where the tile's surface may crystal, indefinite length), discontinuous fiber be aerodynamically heated to 1260°C (Banas, (chopped continuous fiber or staple fiber) and McCormick and Creedon, 1991). Other applidiscontinuous whisker (single crystal, definite cations of these fibers involve reinforcing length). polymer-, glass-, ceramic-, or metal-matrix composites. If the composite is well designed, it will be tougher than the matrix material by 8.2 MANUFACTURE itself. As an example, the fracture toughness of Sic is 1.5 MPa m-2,but 8-15 for Sic reinforced Melt, vapor deposition and chemical processes are used to manufacture fibers and whiskers. with Sic fibers (Richerson, 1992). Fibers and whiskers that are made by melt Fibers can be classified according to strucprocessing include continuous quartz fibers ture, diameter (or cross sectional width) and by drawing from a fused quartz rod, alulength. In general, a material is classified as a mina-silica staple by atomizing a molten fiber if its diameter or cross sectional width is ceramic stream and continuous alumina less than 0.0254 m (0.010 in) and 1ength:diamwhiskers by drawing from a melt. Vapor depoeter ratio is greater than 1O:l.Most commercial sition processes include chemical vapor fibers meet these requirements easily. A fiber is deposition of silicon carbide or boron onto a called a whisker if its microstructure is predominantly single crystal. Fibers or whiskers tungsten core to make continuous fibers and can be subclassified as continuous or discon- the vapor-liquid-solid (VLS) process to make discontinuous whiskers. Chemical processing called sol-gel technology is used to create alumina-based continuous fiber. Other chemical Handbook of Composites. Edited by S.T. Peters. Published processes include the creation of polymer in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
8.1 INTRODUCTION
Continuous fibers precursor fiber that is later pyrolyzed into silicon, carbon and nitrogen-based continuous fiber and the reaction between silica and carbon to make discontinuous whiskers. The manufacturing processes are summarized in Table 8.1. 8.3 CONTINUOUS FIBERS
8.3.1 HIGH SILICA FIBERS
Amorphous high silica fibers can be manufactured by leaching sodium oxide from sodium silicate glass fibers. The glass fibers, made by using conventional glass fiber production techniques, are composed primarily of 74.5%)silica and 24.2% sodium oxide. The fibers are held in a perforated basket, which is suspended within a cylindrical tank. A solution of HC1 and water is pumped from the bottom of the tank up to and over the glass fibers for several hours. The HCl/water/fiber mixture is stirred periodically. The HCl and water are drained away, the fibers rinsed and a second leaching and rinsing cycle is performed. The fibers are dried at a low temperature (Price and Kielmeyer, 1980).Other techniques used to make high silica fibers include spinning from solutions and spinning from a gel (Cooke, 1991).
157
fibers with a minimum of surface flaws, the rod’s surface should be flame polished with an oxy-hydrogen flame, as is done in optical fiber manufacturing (Blyler and DiMarcello, 1991). 8.3.3 CONTINUOUS FIBERS BY CHEMICAL VAPOR DEPOSITION
Producing silicon carbide fibers by chemical vapor deposition (CVD) begins by pulling a melt-spun carbon fiber through a mercury seal and into a tubular reactor. The carbon fiber is heated by coupling the mercury seals on the top and bottom of the reactor to a source of electricity. The melt-spun carbon fiber is called a substrate or a core because silicon carbide is vapor deposited onto it. The deposited silicon carbide is produced by the reaction of a gas mixture of silanes and hydrogen that enters the reactor through a port. Details of producing silicon fibers by CVD are found in Wawner, Jr. (1988) and Textron Specialty Materials (1993b). Producing boron fibers by CVD is a similar process (Tsirlin, 1985; Wawner, Jr., 1988; DeBolt, 1982). 8.3.4 PYROLYSIS OF POLYMERIC PRECURSOR FIBER
Several ceramic fibers composed of various combinations of silicon, carbon, nitrogen, 8.3.2 DRAWING CONTINUOUS FUSED boron and titanium have been produced by QUARTZ FIBERS the pyrolysis of polymeric precursor fiber. The Amorphous fused quartz fibers are made from process consists of four major steps: synthesize preforms of fused quartz rods. The end of a a spinnable polymer, spin the polymer into fused quartz rod is heated with an oxy-hydro- precursor fiber, cross-link (cure) the precursor gen flame and the fiber is drawn, sized and fiber and pyrolyze the cured precursor fiber either taken up on a spool or combined with into ceramic fiber. other simultaneously drawn fibers to form a The synthesis of a polymeric precursor is strand. Fused quartz rods are made by fusing known as Preceramic Polymer Chemistry. chemically purified ground quartz crystal or Many spinnable polymers have been synthequartz sand. Fused quartz rods are classified sized, including polycarbosilane as a as Type I or Type I1 vitreous silica depending precursor for silicon carbide fiber (Yajima, on the fusion technique. Type I signifies that 1985),polytitanocarbosilane as a precursor for fusing was done in an electrically heated cru- Si-Ti-C-0 fiber (Yamamura et al., 1988) and cible, while Type I1 implies fusing with an hydridopolysilazane as a precursor for silicon oxy-hydrogen flame. To obtain fused quartz carbonitride fiber (LeGrow et al., 1987). In
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operation. They claimed that the fiber properties were improved due to uniform heating and even fiber tensioning (Koba et al., 1989). Sumitomo Chemical Co invented a fiber with 85% gamma or delta-alumina and 15%silica by calcining dry-spun fibers. The spinnable mixture was formed by mixing an aluminum compound in a solvent and adding water to get polyaluminoxane. A silica containing compound, an organic polymer to improve spinnability and a compound such as lithium to improve fiber properties were added. The solvent was removed by distillation. The mixture is spun into fiber and the fiber is calcined (Kadokura, Harakawa and Matumoto, 1991). Minnesota Mining and Mfg. Co. (3M) makes fibers composed of alumina, boria and 8.3.5 ALUMINA BASED FIBER silica by drying and sintering fiber that is spun E.I. du Pont de Nemours patented a process to from concentrated aqueous solutions. The make continuous alumina fibers that involved aqueous solution can be composed of an aluspinning a viscous slurry. One slurry composi- mina and boria precursor such as a basic tion was prepared by dissolving an alumina aluminum acetate, silica sol and water. The precursor such as aluminum chlorohydroxide 14%boria in 3M’s Nextel 312 fiber mhibits the in water and then adding solid alumina parti- formation of alumina crystals and aluminum cles. An organic polymer might be added to borosilicate is the predominant crystalline the slurry to aid its spinnability by modifying species. The 2% boria in Nextel 440, however, its viscosity. The viscous slurry was extruded is not enough to inhibit alumina, with etathrough a spinneret and the green (unfired) alumina being the predominant crystalline fibers were wound onto a bobbin. The fibers species (Sowman, 1988). were fired at a temperature below the sintering temperature (calcining) to remove water 8.4 STAPLE and volatile material. The calcined fibers were sintered (densified) to remove porosity and The process to make alumina staple is similar convert the alumina precursor to the stable to the processes that are used in making alumina-based continuous fibers. An aqueous alpha-alumina structure. This slurry-spinning process was advanced and viscous solution of an alumina precursor, by Mitsui Mining Co. where the fiber was spun, silica and an organic polymer is extruded dried, calcined and sintered in a continuous through apertures into a high velocity gas stream, creating green (unfired) staple. The green staple is sintered. The silica is added to toughen the resultant staple by inhibiting aluFootnotes for Table 8.1 mina crystal growth while the green staple is a Most of the trade names are registered trademarks. Each being sintered (Bunsell et al., 1988). trademark belongs to the manufacturer listed to its left. Imperial Chemical Industries (ICI) has a All data on composition for the continuous and staple commercial alumina staple called Saffil. Saffil fibers, except for Tyranno Fiber, are from the manufacturers’ product literature. Tyranno Fiber is 3 ym in diameter and its length is 1-5 cm in composition is from Weddell (1990). the mat form. Saffil comes in six grades ranging
160 Boron, high silica, quartz and ceramicfibers from the catalytic grade with an eta-alumina the end of a polycrystalline feed rod is melted structure, to an aerospace grade with a delta + with a CO, laser, a grain free seed crystal is alpha-alumina or an alpha + delta-alumina + brought in contact with the melt and a whisker is drawn. The LHFZ process can be used to mullite structure. produce whiskers with diameters as low as Alumina-silica staple, also called alumi10 pm or as high as 10 pm (Haggerty,Wills and nosilicate staple, is made by atomizing a Sheehan, 1991). molten ceramic stream. The raw material for the melt is a clay mineral or a blend of alumina and silica. The melt is atomized by pouring it 8.5.2 DISCONTINUOUS WHISKERS BY THE VLS into the path of a high velocity air stream or by PROCESS feeding it onto a rotating wheel that throws out the melt by centrifugal force. The former is VLS stands for vapor feed gases-liquid catacalled the blowing process, while the latter is lyst-solid crystalline whisker growth. In a called centrifugal spinning. One commercial particular VLS process to make single crystal alumina-silica staple called Fiberfrax ranges Sic whiskers, uniformly sized catalyst particles in length from 70-1000 pm and is made by The are applied to carbon substrates. The carbon substrates and catalyst are placed within a Carborundum Co. growth chamber. Also within the chambers are Si0 generators, made by impregnating porous brick with carbon and SiO,. The chamber is 8.5 WHISKERS placed within a furnace and heated to a growth 8.5.1 CONTINUOUS WHISKERS temperature of 1300-1450°C. During t h s time, Single crystal continuous alumina whiskers can gases such as CO, CH,, N, and H, are fed into be grown from a pure melt by using a modifi- the chamber. The catalyst particles melt and cation of the Czochralski method. The modified become supersaturated with material supplied method is still classified as a moving crystal from the vapor and carbon from the substrate. techruque, but, unlike Czochralslu’s technique, Solid Sic precipitates from the liquid catalyst uses a die to shape the growing whisker. A cap- onto the growth substrate. As precipitation conillary tube draws the melt up to and over a die. tinues, the whisker grows, lifting the catalyst The melt flows to the die’s vertical edges and ball from the substrate (Petrovic and Hurley stops; so the die’s vertical edges control the 1990; Milewski et al., 1985). The time necessary to produce Sic whiskers shape of the subsequently drawn whisker. A grain free seed crystal is brought in contact with by the VLS process can be reduced by underthe melt surface and a whisker is drawn. A lin- cooling or pre-alloying the catalyst balls. After ear temperature distribution is maintained reaching the growth temperature, the catalyst along the axis of the growing crystal to avoid balls are undercooled by 150”C,initiating predefects by lowering thermal stresses (LaBelle, cipitation. The chamber is brought back up to Jr, 1980; Antonov, 1990). Saphikon Inc. grows the growth temperature and held there until whiskers in many shapes using a similar whisker growth is completed. Another techprocess that they call EFG (edge-defined,film- nique to reduce the growth time is to pre-alloy fed growth). One of the shapes is a hollow tube the catalyst particles with carbon and silicon that may be used as a waveguide (Harrington so that whisker growth occurs immediately upon reaching the growth temperature and Gregory 1990). Another moving crystal technique to make (Shalek, 1987; Shalek, Katz and Hurley, 1988). continuous alumina whiskers is called the Other whiskers produced by the VLS laser heated floating zone directional solidifi- process include A1,0,, boron, MgO and Si cation process (LHFZ). In the LHFZ process,
Forms 161 (Wagner, 1970); Tic by using TiC1, and CH, as the primary feed gases and a nickel catalyst (Pearson and Easley, 1992; Narasimhan and Bhat, 1992);Si,N, by using Si,Cl, and NH, as the primary feed gases and an iron catalyst (Motojima et al., 1989);and TiB, by using TiC1, and BC1, as the primary feed gases and Au or Si as the catalyst (Withers, Loutfy and Lee, 1988). 8.5.3 OTHER DISCONTINUOUS WHISKER GROWTH PROCESSES
An early vapor phase growth process to make alumina whiskers involved hydrogen reduction. Wet hydrogen was passed over heated aluminum powder. Whisker lengths and diameters were from 1-30 mm and 3-50 pm, respectively. Hydrogen reduction of methyltrichlorosilane was used to produce p-Sic whiskers (Milewski, 1991).Other vapor phase growth techniques are evaporation-condensation and vapor phase reaction. These and other growth techniques, are discussed in Bracke, Schurmans and Verhoest (1984). Some commercial processes to make Sic whiskers use a chemical reaction between silica and carbon powders. Tokai Carbon manufactures p-Sic whiskers by heating a mix of pulverized silica gel, carbon black and a metal catalyst. The mix is heated in a nonoxidizing atmosphere at 1300-1700°C and the whiskers grow within the pores of the mix (Yamamoto, 1985).The whiskers are under 1.0 pm in diameter and under 300 pm in length. Other companies have produced Sic whiskers by chemical reactions: the J.M. Huber Corp. passed perforated trays filled with ground rice hulls over heated carbon pellets and through a horizontal furnace. Ground rice hulls were used as the reactant material because they contain both silica and carbon (Tanaka, Kawabe and Kobune, 1986). Silicon nitride whiskers have been produced by heating a silica and carbon mix or rice hulls in a nitrogen atmosphere (nitriding), or by heating silica powder in a gas mixture of ammonia gas and hydrocarbon gas.
8.6 PROPERTIES
Boron, h g h silica, quartz and ceramic fibers have a high thermal and chemical stability relative to organic fibers. Maximum use temperatures of boron, h g h silica, quartz and ceramic fibers, except for boron fiber in an oxidizing atmosphere, all are over 1000°C. These fibers have a good balance of specific strength (tensile strength/density ratio) and specific modulus (elastic modulus/density ratio) compared with other materials. Tsirlin’s (1985) graphs of specific strength against specific modulus for boron or Sic fibers with cores illustrate this point. In Table 8.2 note the low coefficient of thermal expansion for quartz fibers, which allows them to be used in parts that are subjected to high temperatures, such as in high speed printed circuit boards. 8.7 FORMS
The form that a fiber can take depends on its length, diameter and mechanical properties, such as elastic modulus and tensile strength. In general, discontinuous whiskers are too short to be spun into yarn, however, VLS whiskers about 75-100 mm long might be spun into staple yarn (KO, 1989).A fiber’s flexibility is related to its diameter, elastic modulus and tensile strength. A measure of flexibility is l / [ d ( E / o- l)]where d = diameter, E = elastic (Young’s)modulus and cs = tensile strength. A greater value indicates a greater flexibility. Obviously, the greater a fiber’s flexibility, the easier it is to twist, braid and weave into complex shapes without breakage. As an example, l / [ d ( E / o - l)]is about 126 for continuous Sic fiber with a core (d = 80 pm), 533 for continuous alumina fiber, 1298 for continuous Sic fiber (d = 12 pm) and 10 101 for continuous fused quartz fiber. The Sic fiber with a core, having a lower flexibility relative to other fibers, is generally used for unidirectional reinforcement. The fiber might also be used in a material-geometric hybrid to axially reinforce a woven preform of finer
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Footnotes for Table 8.2 All values, except those for silicon carbide whiskers, are from the manufacturers’ product literature. Test methods may varv between manufacturers. All values are averages, are for reference only and are not to be taken as a specification. Most fibers are under continual development. Contact the manufacturers for the most up to date property values. DNA = Data Not Available. The silicon carbide whisker values are from Petrovic and Hurley (1990). This value in metric units was converted from a value that was given in English units. Most of the tradenames are registered trademarks. Each trademark belongs to the manufacturer listed to its left
and two diagonal directions. Figure 8.1 shows the 3D and 5D fabrics. See Ngai (1990) for diagrams of 4D, 7D and 11D fabrics. Continuous boron fibers are supplied in the form of borodepoxy prepreg tape and boron/aluminum preform sheets. The boron/aluminum preform sheets are made by resin bonding the boron fibers to the side of aluminum foil. The fibers can be aligned in any direction. These preform sheets are used in solid state (low temperature, high pressure) diffusion bonding processes (TextronSpecialty Materials, 1993a). Continuous Sic fibers with a core are formed into plasma-sprayed aluminum or titanium preform sheets for hot molding (a low pressure, hot pressing process) and into ’woven fabric’ for investment casting or diffusion bonding. In diffusion bonding, the fabric is placed between two metal foils prior to consolidation. The ’woven fabric’ is similar to nonwoven roving because the fibers are aligned in one direction and held together by a cross-weave of metallic ribbons (Textron Specialty Materials, 199313; Mittnick 1990). Continuous high silica fibers are processed into a variety of forms such as cloth, tape, sleeving, yarn, cordage, rope, nonwoven mat, blanket, felt, rigid tile and sewing thread. Alumina-silica staple fibers are made into paper, rigid board and cylinders, other rigid three-dimensional shapes and into most of the high silica fiber forms. Alumina staple fibers are processed into paper, mat, felt, blanket, rigid board and cylinders and other rigid three-dimensional shapes.
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Table 8.3 provides a sampling of the many applications for boron, high silica, fused quartz and ceramic fibers. For other applications, see Bracke, Schurmans and Verhoest (1984), Mortensen and Koczak (1993), Schwartz (1992) and the other chapters within this handbook.
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166 Boron, high silica, quartz and ceramic fibers 8.8.1 NONCOMPOSITES
Alumina and alumina-silica staple fibers and the continuous high silica fibers are used in high temperature applications in many industries. Forms of these fibers are used for thermal insulation, personnel protection and gasketing. See Carborundum (1991) for the many applications of high silica fibers and Carborundum (1989) for the many applications of alumina-silica fibers. 8.8.2 COMPOSITES
Fibers and whiskers have been used to reinforce polymers, epoxies, glasses, metals, or ceramics. The reinforced materials are called composites in general and, more specifically, polymer, epoxy, glass, metal, or ceramic matrix composites. The reinforcements are either discontinuous whiskers or continuous fibers for the applications in Table 8.3, but staple fiber and continuous whisker may be used as a reinforcement material as well. Usually the fiber or whisker is coated to cover its surface flaws, provide a diffusional barrier between it and the matrix material and protect the fiber or whisker from oxidation. In addition, the coating provides a weak interface that deflects cracks propagating through the matrix in ceramic matrix composites REFERENCES
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167
Price, G.B. and Kielmeyer, W.H. 1980. Method for Making I-ligh Purity, Devitrification Resistant, Amorphous Silica Fibers. US Patent 4 200 485. Apr. 29, 1980. Assigned to John-Manville Gorp. Richerson, D.W. 1992. Modern Ceruniic Engineering, 2nd edn. New York: Marcel Dekker, Inc. Schwartz, M.M. 1992. Composite Materials I$andbook, 2nd edn. ed. M.M. Schwartz. New York: McGr a w-Hill , Inc . Shalek, P.D. 1987. Process for Growing Silicon Carbide Whiskers by Undercooling. US Patent 4 702 901. Oct. 27, 1987. Assigned to The United States of America. Shalek, P.D., Katz, J.D., Hurley, G.F. 1988. Prealloyed Catalyst for Growing Silicon Carbide Whiskers. US Patent 4 789 537. Dec. 6, 1988. Assigned to The United States of America. Schramm, Dale E. 1988. Process for Producing Silicon Carbide Whiskers. US Patent 4 789 536. Dec. 6, 1988: Assigned to J.M. Huber Corp. Sowman, H.G. 1988. Alumina-Boria-Silica Ceramic Fibers from the Sol-Gel Process. In Soi-Cel Technology for 7 h i n Films. Fibers. Prefornzs. Electronics. und Specialty Shapes, ed. Lisa C . Klein, pp. 162-83. New Jersey: Noyes Publications. 'Takeda, M. rt al. 1992. Thermal Stability of the Low Oxygen Silicon Carbide Fiber Derived From Polycarbosilane. Ceram. Eng. Sci. Proc. 13 (7--8): 209-1 7. Tanaka, M., Kawabe, T. and Kobune, M. 1986. Method of Manufacturing Crystalline Silicon Carbide Employing Acid Pretreated Rice Husks. US Patent 4 591 492. May 27, 1986. Assigned to Tateho Kagaku Kogyo Kabushiki. Tech Trends 1990. Ceramic Matrix Composites: Technology and fndustriul Applications. Paris: lnnovation 128. 'Textron Specialty Materials 1993a. Boron Aluminum Preform Sheets. Textron Specialty Materials. Lowell, MA, USA. Textron Specialty Materials 1993b. Continuous Silicon Carbide Metal Matrix Composites. Textron Specialty Materials. Lowell, MA, USA. Tricoles, G.P. 1988. Radome Electromagnetic Design. In Antenna Handbook, ed. Y.T. Lo and S.W. Lee, pp. 31-1 to 31-31. New York: Van Nostrand Reinhold. Tsirlin, A.M. 1985. Boron Filaments. In Strong Fibres, ed. W. Watt and B.V. Perov, pp.155-99. Vol.1 of Handbook of Composites, Series ed. A. Kelly and Yu. N. Rabotnov. Amsterdam: Elsevier Science Publishers R.V.
168 Boron, high silica, quartz and ceramicfibers Wagner, R.S. 1970. VLS Mechanism of Crystal Growth. In Whisker Technology. ed. Albert P. Levitt, pp. 47-119. New York John Wiley & Sons. Wawner, F.E. 1988. Boron and Silicon Carbide/Carbon Fibers. In Fibre Reinforcements for Composite Materials, ed. A.R. Bunsell, pp. 371425. Vol. 2 of Composite Materials Series, Series ed. R.B. Pipes. Amsterdam: Elsevier Science Publishers B.V. Weddell, J.K. 1990. Continuous Ceramic Fibers. 1. Text. Inst. 81 (4): 333-59. Withers, J.C., Loutfy R.O. and Lee, C.T. 1988. Process to Produce Titanium Diboride Whiskers as Reinforcement for Metal and Ceramic Composites. NSF Grant ISI-8760300. National Science Foundation, Washington D.C. October 1988.
Xiao, H., Ai, X and Yang, H.S. 1993. Effect of whisker orientation on toughening behavior and cutting performance of SiCW-A1203 composite. Mat. Sci. Technol. Vol. 9:21-25. Yajima, S. 1985. Silicon Carbide Fibers. In Strong Fibres, ed. W. Watt and B.V. Perov, pp. 201-37. Vol.1 of Handbook of Composites, Series ed. A. Kelly and Yu. N. Rabotnov. Amsterdam: Elsevier Science Publishers B.V. Yamamoto, A. 1985. Process for Preparing Silicon Carbide Whiskers. US Patent 4 500 504. Feb. 19, 1985. Assigned to Tokai Carbon Co Ltd. Yamamura,T., et al. 1988. Development of a New Continuous Si-Ti-C-0 Fibre Using an Organometallic Polymer Precursor. 1.Mater. Sci. 23: 2589-94.
CARBON FIBERS
9
Khalid Lafdi and Maurice A. Wright
9.1 INTRODUCTION
microstructural features developed in the spinCarbon fibers exhibit truly outstanding prop- ning process that also influence the stiffness, erties. As shown in Table 9.1, their strength, 0, strength and the thermal and electrical propercompetes with the strongest steels; they can ties of pitch can be optimized. Nevertheless, have stiffness, E , greater than any metal, the tensile strength of PAN-based fibers has ceramic or polymer; and they can exhibit ther- always been superior to pitch-based fibers; since PAN fibers were developed before pitch mal and electrical conductivities that greatly fibers, most structural materials and compoexceed those of competing materials. If the nents use PAN-based fibers. strength or stiffness values are divided by the The most important mechanical and physilow density, p, 1800-2200 kg m-3, then their cal properties exhibited by carbon fibers are the huge specific properties (o/p, E/p) make this elastic modulus, tensile strength and the electriclass of materials quite unique. cal and thermal conductivities. These All carbon fibers sold commercially are fabproperties are sensitive to the crystallite size ricated from polyacrylonitrile (PAN) or from a and perfection of the graphene layers develcoal, petroleum or synthetic pitch. PAN-based oped within the carbon fiber and depend for fibers are produced from a solubilized mixture the most part on the degree of molecular alignthat is wet or dry spun to produce a fiber, ment with respect to the fiber axis. Growth and ostensibly for use in the textile industry. This alignment of such layers occur within the prefiber is stabilized and carbonized to produce a cursor and within the solid carbon fiber when it carbon fiber. Aerospace grade material can be is heated to high temperature. However, it can obtained in tows that contain between 3000 be argued that the most extensive structural and 12 000 fibers. Lower performance materirearrangement occurs when the basic structural als are usually formed using larger tows that units (BSUs) present in the original precursor contain up to 320 000 fibers. PAN-based carare large and plate-like so that the shear stresses bon fibers are cheaper when produced from generated during spinning can align these large larger tows. areas more easily. Heating the fiber to high temPitch fibers are melt spun products obtained in small tow sizes varying from 2000 to 4000 perature after spinning relaxes the spinning stresses and allows the oriented regions to fibers. They are usually of larger diameter grow. Compared to PAN, the basic structural (- 10-15 pm) than fibers formed from PAN. units in an original mesophasic pitch are much The spinning process is controlled by the larger in area and length and are not twisted to pitch-based carbon fiber producer; thus, the same degree. The resulting ease of graphitization and alignment of the graphene layers and the development of large crystallites, proHandbook of Composites. Edited by S.T. Peters. Published duces a large elastic modulus and electrical and in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
-
170 Cnrbon fibers thermal conductivity parallel to the fiber axes. Unfortunately, large graphitized regions tend to produce high local stress concentrations, especially if they are misaligned (Reynolds and Sharp, 1974; Reynolds and Moreton, 1980). 'rhus, pitch-based fibers tend to exhbit high modulus and high electrical and thermal conductivities, but low strength. PAN-based fibers tend to be of intermediate modulus and relatively high strength. The mechanical properties of structural metals generally vary in direct proportion to density. Thus, as shown in Table 9.1, the specific
properties (o/p, E/p) tend to remain relatively constant; structures designed to resist a given set of loads tend to weigh the same irrespective of whether they are made from aluminum, titanium or steel. These mechanical property/ density relationships are not followed by carbon fibers; so, compared with metals, similar structures made from carbon fiber reinforced composites will be lighter. The most successful use of carbon fiber reinforced materials has been for military purposes, especially aeronautical structures (Schwartz, 1984; Hadcock, 1982). Composite
Table 9.1 Physical and mechanical properties of typical carbon fibers -
Siipplier/Fiber name
Tensile strnzgth, a (GPd
Elastic rnoduliis, E (CPal
Fracture elongation,
Density, dp. p ( k g / ~ ~ IOp ) rn? s-:
E/P > I O h rn2s-'
%
Pitch Type Amoco /Thornel
PlOO P120
2.2 2.4
690 830
0.3 0.3
-2150 -2150"
1.02 1.11
321 386
Petoca /Carbonic
HM50 HM60 HM80
2.8 3.0 3.5
490 590 790
0.6
-2000"
1.40
245
0.5 0.4
-2000"
1.75
395
T-40
3.45 2.9 5.65
231 390 290
NA NA NA
1760 1810 1810
1.96 1.60 3.12
131 215 160
Akzo / Fortafil
F-~(c) F-3(c)
2.76 3.80
345 227
NA NA
1800 1700
1.53 2.24
192 133
RK Fibers/RK
RK30 RK25
~3.0 >2.5
-230 -230
NA NA
1780 1780
1.68 1.40
129 129
Toray/Torayca
M46J T300
436 230 294
0.5 1.2 2.0
1750
1.71
143
T800
4.2 3.5 5.5
ST-1 ST-2 ST-3
3.6 4.0 4.4
240 240 240
1.5 1.7 1.8
-1800b -1800b -1800b
2.00 2.20 2.38
160 141 133
0.172 0.324 0.414
73 110 199
NA NA NA
2720 4500 7860
0.063 0.072 0.052
27 24 25
PAN Type Amoco/Thornel
T-300
T-50
Toho/Besfight
Metals Aluminum' Titanium' Steel'
"Estimated from Tanabe et al. 1987; hEstimated from Johnson 1987; 'Eshbach and Souders, 1974; NA, Not Available.
Overview 171 use in civilian aviation has been more limited, largely because of cost. However, in 1985 the European Airbus consortia used carbon/epoxy vertical stabilizers on their A310 and, since 1993, delivered their A340 and A330 models equipped with composite tail sections, floor panels, landing gear doors and carbon-carbon brakes. The new Boeing 777 and the projected McDonnell-Douglas MD-12 contains similar composite structures (High-Performance Composites, 1994).
a precursor material is prepared, spun into a fibrous shape, stabilized to change it from a thermoplastic to a thermoset, then heated until all unwanted elements are expelled. Depending on the final heat treatment temperature, a fiber is produced that is primarily carbon. A strikingly similar production technique is used to produce either type of fiber; however, the initial pretreatment and the chemical reactions that occur within either PAN or pitch during stabilization and carbonization are markedly different. These differences are discussed in the following sections since they determine the ultimate properties and the cost of the carbon fibers that are produced.
9.2 OVERVIEW
Ln this chapter we have chosen to combine, rather than separate, the discussion of the preparation of carbon fibers from PAN or from pitch. This style of presentation is attempted since the methods used to fabricate carbon fibers from either of these precursor materials seem to be almost identical, as shown in Fig. 9.1; for example, both methods involve the preparation, spinning and subsequent thermal degradation of organic precursors. Essentially,
9.2.1 POLYACRYLONITRILE
The chemical composition of PAN-based precursors tends to be proprietary; however, in general it consists of small diameter linear molecules that are made up from nitrogen, hydrogen and carbon (Fig. 9.2). Spinning tends to orient these molecules parallel to the
Pitch precursor
Polyacrylonitrile
I
Particulate removal
I
Pretreatment (chemical, thermal, mechanical)
I
Wet or dry spinning
I
Fiber spinning
Solvent extraction
I
Stabilization at 200-260°C
I
Carbonization to 1500°C
I
Graphitization to 2500°C Fig. 9.1 Schematic illustration of process to produce carbon fibers from polyacrylonitrile and pitch.
172 Carbon fibers
\
CH
CH
CH
CH
C
C
C
C
C
n
NN
/CH2\
a
NN
/CH2\
/
,CH2\ CH
a
/CH2\
a
n
NN
NN
NN
Fig. 9.2 Structure of the ideal PAN molecule (Eggs, 1976).
fiber axis, but they continue to be randomly oriented transverse to that direction. Thus, a fiber with a twisted fibrillar structure is produced. This fiber is supplied by the textile industry to the carbon fiber fabricator who stabilizes it under tension before converting it into a carbon fiber using a controlled heat treatment process. Apart from the obvious generation of the fibrous shape, the importance of the spinning process to the fabrication Initiation
of carbon fibers from PAN is relatively minor; more important is the chemical makeup of the PAN and the presence of small amounts of other constituents that influence the complex chemical reactions that occur during stabilization and carbonization. Stabilization involves cyclization of the oriented molecules and results in the release of most of the hydrogen and part of the nitrogen as NH, and other nitrogen compounds (Fig. 9.3). The role of the Cyclization
Abnormal structures end structures impurities N
H Abstraction
n = 0, 1 , 2 ... m = 0,1,2.
2. Transfer 1.Cyclization reactions
scission
I \
N intra- or intermolecular transfer
I
re-initiation
Fig. 9.3 Release of HCN and NH, during cyclization of PAN fiber (Grassia and McGuchan, 1971b).
Pitch precursor treatments 173 retained nitrogen is very important to both the crosslinking process and to the development of optimum properties during carbonization. 9.2.2 PITCH
A pitch precursor taken from a petroleum or coal tar feedstock initially contains individual molecules that exhibit appreciably different molecular weights. These untreated precursors have been used to produce fibers; however, they are isotropic and exhibit relatively poor mechanical and physical properties. Conversely, the carbon fiber producer can pretreat the pitch to develop a continuous anisotropic phase (similar to a mesophasic liquid crystal) or a two phase mixture that becomes highly oriented during the subsequent spinning process. In contrast to the chemical changes occurring in PAN, physical changes are responsible for the final properties of pitch-based carbon fibers. Essentially, the precursor isotropic pitch is pretreated to produce a two phase mixture that is predominantly anisotropic. During spinning and drawdown, this mixture is very strongly oriented both parallel and transverse to the fiber axis. Oxygen added during stabilization tends to crosslink these large molecules in a simple way before being released on carbonization as CO, CO, and H,O. More recent interest has centered on producing fibers from synthetic pitches. These require no extensive pretreatment and they stabilize faster at a given temperature. Fibers made from these pitches should cost less to produce. Precursor materials that have been used to produce carbon fibers include: polyamides, polyesters, polyvinyl alcohol, polyvinylidene chloride, poly-p-phenylene, phenolic, napthalene, naphthalene-phenanthrene, alkylbenzenes, rayon, polyacrylonitrile (PAN) and various petroleum, coal tar and synthetic pitches (Ezekiel, 1969; US Patent 3 533 741, 1970; Shindo, Nakanishi and Soma, 1969; Boncher, Cooper and Everett, 1970; French Patent 1535 800, 1968; Kawamura and Jenkins, 1970;
Mochida et al., 1988; Lewis and Nazem, 1987a). Today, fibers are produced commercially from rayon, PAN and the various pitches. However, the process to produce fibers from rayon is very expensive because it involves stretching at very high temperatures and the yield of carbon after carbonization is small. Rayon-based fibers are therefore fabricated in such small amounts that they are really of no commercial importance. Additional information can be obtained from: Yanagida et al., 1991; Bacon, 1973; US Patent 3107152, 1963; Yoneshoga and Teranishi,l970. 9.3 PITCH PRECURSOR TREATMENTS
A typical pitch precursor material is obtained from either the distillation products involved in the chemical treatment of decanted oil or as a by-product of the production of metallurgical coke from coal. In a conventional as-received pitch, basic structural units (BSUs) are already present to a degree that pitch can be considered a random suspension of highly aromatic molecules similar to coronene (molecular weights between 600 and 900) surrounded by a liquid of smaller molecular weight. Heating such a system initially reduces the viscosity. Eventually, however, the viscosity increases as the BSUs grow and coalesce to form larger entities (molecular weights between 1000 and 4000) with a specific long range anisotropy called 'mesophase' (Brooks and Taylor, 1965). Continued heating eventually causes an inversion in which the mesophase spheres become the continuous anisotropic phase within which are suspended spheres of the isotropic low molecular weight material. The rate of viscosity increase is very slow at low temperatures, but accelerates as the temperature is increased. At temperatures greater than 350°C, the pitch begins to form coke by a process of thermal degradation and gas evolution. The variation in viscosity with temperature for numerous pitch fractions has been documented (Bathia, Fitzer and Kompalik, 1984) and is shown in Fig. 9.4. As
174 Carbon fibers
60
I-
lS
CTP A240
1:1
1:4
I I I
45
I
-
I
I I I
v)
B
>. c .v) 0
I
I I
30-
.-> v)
I
U
I
C
EQ
P
I 1s-
\
50
100
150
200
250
300
350
b 00
1so
500
550
Temperature OC Fig. 9.4 The variations of apparent viscositv w ~ t l temperature i o f T ariotii pitch fractions (Batha, Fitzer and Kompalik, 1984)
can be noted, at a given temperature, the viscosity of any pitch is greater the more anisotropic phase it contains. By cooling after partial reaction, it is possible to produce a two phase pitch with a viscosity appropriate for spinning or infiltration. A number of experiments have been carried out designed to accelerate the process of producing two phase pitch mixtures. These involved heating, stirring, bubbling an inert gas through the liquid (sparging), or the combination of stirring and sparging (US Patent 3 629 379, 1971; US Patent 3 919 383, 1975; US Patent 3 974 264, 1976; US Patent 4017327, 1977). In this way, an appropriate anisotropic concentration can be produced in much shorter time periods (hours instead of days) with improved spinning characteristics. This latter characteristic seeins to be associated with the smaller molecular size existing within the two phase mixtures formed by stirring and sparging and the smaller variation of molecular size.
A 1980 patent (US Patent 4 208 267, 1980) discussed a different method of producing isotropic/anisotropic pitch mixtures. The method essentially consists of dissolving part of the original isotropic pitch in an organic solvent such as benzene, toluene, xylene, etc. The material that is insoluble can then be converted, by heating, into a material that is greater than 75% anisotropic. The efficiency of this process is quite poor however, since only a \Tery small amount of this 'neomesophase' can be produced from a given pitch. For instance, using Ashland A260 pitch, about 75-90% of the initial pitch will dissolve. Using Ashland A240,80-905:1 of the pitch disso1v es. Variations of the gas-sparge process fpossibly associated with a chemical fractionation) can be made to change the characteristics of a resultant pitch. Lafdi, Bonnamy and Oberlin (1991a; 1991b; 1991c; 1992) and Lafdi and Oberlin (1994a;1994h) have indicated that
Spinning conditions 175 some pitches exposed to a nitrogen sparge at atmospheric pressure or a hydrogenation treatment at high pressure produce a continuous strongly anisotropic material that contains small particles of a weakly anisotropic material. During spinning, the second phase becomes completely absorbed (or transformed) to produce a uniformly anisotropic fiber. They believe that sparging disturbs the formation of an anisotropic phase that exhibits large differences in molecular weight. Indeed, they suggest that the spheres of anisotropic material contain BSUs that are only weakly associated. In contrast to the strong molecular orientation exhibited by the Brooks and Taylor type of ’mesophase’ (Brooks and Taylor, 1968), the common orientation of this new anisotropic material results from the statistical orientation of small units. The pitch then behaves as a two component gel which exhibits a long range anisotropy in the bulk. The short range change in orientation of the carbon units produces sharp changes in orientation similar to grain boundaries. Such regions produce a zig-zag nanotexture in the resultant carbon fiber that prevents inter-sheet gliding and provides a crack inhibiting function that contributes to relatively high ultimate strength values.
9.4 SPINNING CONDITIONS
9.4.1 POLYACRYLONITRILE
Dry and wet spinning of polyacrylonitrileprecursors have been used. In the dry process the polymer is solubilized and spun into a current of hot air that removes the solvent. Unfortunately, solvent removal tends to be quite rapid and can cause the outer portion of the fiber to solidify before the solvent can diffuse from the fiber’s center. The large diffusion gradient that develops can seriously affect the final shape of the fiber (Edie and Diefendorf, 1993). The more common wet spinning method involves solubilizing the polymer with a polar solvent such as dimethyl
acetamide before extruding it into a ’coagulation’ bath through a spinneret. In the wet-spinning process, the fiber is solidified by using a coagulant (such as ethylene glycol) which extracts the solvent from the polymeric fiber. In a manner similar to the dry spinning process, the rate at which the solvent is extracted from the polymer as it passes through the coagulation bath governs the final shape of the fiber. The temperature, concentration and circulation rate of the fluid in the coagulation bath are known to affect the structure and hence the physical and mechanical properties of as-spun fiber. Many companies have added a supplemental stage to the spinning process that is designed to reduce the water content. This additional step tends to increase the molecular orientation within the fiber (US Patent 3846833, 1975; US Patent 3 841 079,1974). A typical acrylonitrile-based precursor contains several percent of various co-monomers such as methyl acrylate or vinyl acetate which improve the precursor’s spinnability or fabric properties. Though not added to aid carbonization specifically, they have been found to influence the properties of the resulting carbon fiber. Many modified PAN polymers such as acrylonitrile-hydroxyethylene, acrylonitrile-vinyl chloride-itaconic acid (French Patent 2 328 723), polyacrylomidoxium (US Patent 3 767 773,1973) have been investigated to obtain a suitable as-spun fiber capable of producing a fiber with a large carbon yield after carbonization. 9.4.2 PITCH According to Singer (US Patent 3919383 (1975)), in order to spin a fiber, pitch must be heated to produce a viscosity between 10 and 200 poise (1-20 Pa s). However, temperatures greater than about 35OOC cannot be used to obtain the required viscosity because thermal decomposition of the pitch will occur. In addition, spinning should be carried out above a minimum temperature of about 200°C since
176 Carbon fibers The spinning of two phase mixtures is not an easy commercial operation since the phases exhibit different viscosities and densities. The strongly anisotropic continuous phase contains within it a less anisotropic (or isotropic) phase which exists as spheres. If the diameter of the spheres is large relative to the spinning orifice, localized weak sections of extrudent can be produced that can break and make it difficult to maintain a continuous fiber thread. In addition, since the stabilization rates of each phase differ, one phase may be over stabilized relative to the other and it becomes difficult to
this determines the maximum temperature that can be used in the subsequent stabilization step. These temperature requirements define a processing window into which suitable pitches must fall. Using literature data (Lewis and Nazem, 1987a;Mochida et al., 1988; Yanagida et al., 1991), White (1992) has shown that the smallest window exists for 100% anisotropic pitches processed from coal or petroleum. A larger window exists for material partially transformed from the same precursors; however, as shown in Fig. 9.5, the largest window exists for synthetic pitches.
Temperature ("C) 450
350
400
250
300
200
Window Boundaries
4 c
h
In (d
4
> + .-
11
Sol poi
I
In
8 In
5
h
In lU" (d
4
>
c .-
lo2
In
0 0 In
5 A 7/////
B
\
I
1.4
l
l
1.5
l
\ Lowest viscosity pitch
TC I
l
1.6
l
1'"'
that can be stabilized
-
High temperature limit, to avoid pitch decomposition t
1.7
l
l
1.8
l
l
1.9
l
l
2.0
l
l
2.1
,
10-3 2.2
Inverse Temperature, I/Tx lo3 (K-l) Fig. 9.5 Processing window for injection of mesophase pitch (White, 1992). A A : The lowest viscosity reported for mesophase pitch prepared from petroleum or coal-tar pitch by pyrolysis to 100% mesophase. BB': The lowest viscosity reported for a petroleum- or coal-tar-based mesophase pitch only partially transformed but with the mesophase acting as the continuous phase. CC': The lowest viscosity reported for a chemically-derived fully-transformed mesophase pitch.
Stabilization of polyacrylonitrile generate optimum properties in the resulting carbon fiber. This fact seems to have been recognized early, since Singer (US Patent 3 919 383, 1975) describes methods to remove the isotropic component from the two phase fiber immediately after spinning. Due to the proprietary nature of the various procedures, little is known of the spinning process itself, however, a number of patents have been awarded (US Patent 4 331 620,1982; Matsumoto 1985; US Patent 4 504 454, 1985). These have been discussed in other publications (Wright, 1989; Wright and Palmer, 1994); most deal with attempts to combat the two phase nature and non-Newtonian behavior by stirring the pitch in the spinneret. Apparently, agitation of mesophase within the spinneret is part of the technique used by the Petoca Oil Company to randomize the cross sectional microstructure and produce the folded graphene layers that are thought responsible for the relatively high strengths exhibited by their ’carbonic’ carbon fiber (Guigon, Oberlin and DesarmotJ984a; 1984b). As seen previously in Table 9.1, for similar modulus materials, the strength of Petoca’s carbonic fibers (HM-series), while still not as good as PAN based fibers (T-series),is appreciably better than other commercially available pitch material (P-series). The flow of pitch through the spinneret (Matsumoto, 1985), the geometry of the die hole (Singer, 1978; Yamada et al., 1984; US Patent 4 504 454, 1985), pitch viscosity, die swell and drawdown are all known to affect the microstructure of the resulting pitch-based fiber. Indeed, the above authors illustrate their comments with fiber microstructures that vary from circumferential through random, radial or radial with a crack (Fig. 9.6). Spinning experiments using 100% ’mesophasic’ synthetic pitches produced from napthalene (Ohtsuka, 1988), napthalene-phenanthrene (Lewis and Nazem, 198%) and alkylbenzenes (Yanagida et al., 1991) have been reported. The viscosity of these materials is appropriate for spinning
177
Fig. 9.6 Typical pitch-based carbon fiber microstructures (Edie, 1990): (a) radial; (b) onionskin; (c) random; (d) flat-layer; (e) radial-folded;(f) radial with wedge.
at temperatures lower than those necessary to spin anisotropic-isotropic mixtures obtained from petroleum or coal tar pitch; thermal decomposition during spinning is therefore not a problem. Stabilization of such pitches must be performed at correspondingly low temperatures; however, when they are compared to petroleum and coal tar mesophasic pitches, stabilization at a given temperature occurs at a significantly faster rate. This feature, when combined with the fact that no pitch pretreatment is necessary, should enable fibers to be produced at significantly lower cost. The result of all spinning processes is a PAN or pitch fiber that can be changed into a carbon fiber by heating at a very slow rate to a temperature of about 1500°C. A higher rate of heating will usually melt the fiber unless it is first converted from a thermoplastic to a thermoset. This conversion or stabilization treatment is done by heating to a relatively low temperature for an extensive time period in an atmosphere containing oxygen (usually air). 9.5 STABILIZATION OF POLYACRYLONITRILE
This process converts the thermoplastic, as-spun polymer into a thermoset that is capable of maintaining its shape during
178 Carbon fibers carbonization (Lavin, 1992; Yooh, Korai and Mochida, 1994).The operation is identical for both PAN and pitch based fibers. However, the stabilization of anisotropic pitches involves simple cross linking of plate-like molecules whereas the stabilization of PAN involves many different chemical reactions. Stabilization of both PAN and pitch is an exothermic process, so great care must be taken to control the rate of reaction and to avoid thermal runaway which melts the fiber and is a fire hazard. Commercial stabilization is carried out by heating the PAN fiber in air between 200 and
CYCL’ZATloN
I
CN
260°C for a period of time that varies between thirty minutes and several hours. During stabilization, several interdependent chemical reactions occur. The reaction that dominates is primarily determined by the chemical composition of the initial precursor, the spinning history, the final composition of the as-spun fiber and the stabilization heating schedule. A PAN polymer mainly consists of acrylonitrile entities -CH,CH(CN)- which are able to cyclize (Johnsonet al., 1972)with the help of an initiator into a presumably linear ’ladder polymer’ similar to that shown in Fig. 9.7. In general, the pendant nitrile groups of PAN
CN CN PAN
Fig. 9.7 The process of PAN stabilization and subsequent carbonization (Fitzer and Heine, 1988).
Stabilization of polyacrylonitrile first become crosslinked to form a ladder polymer. Initiation of this process is catalyzed in some cases by the presence of a small amount of reactive copolymer such as itaconic acid. Oxygen is then incorporated into the ladder polymer under a number of possible schemes which have been described by Watt and Johnson (1975) and by Clarke and Bailey (1973) and are shown in Fig. 9.8. Cyclization and stabilization induce tremendous shrinkage into the polymer. Longitudinal shrinkage is resisted mechanically; however, the diameter of the fiber is allowed to decrease.
A balance should be kept during stabilization as to hydrogenation degree. A large hydrogen content can result in a small N/C ratio which increases the temperature at which the local molecular ordering occurs. Conversely, increasing the available oxygen decreases the size of and the temperature at which the units of local molecular order (LMOs)are formed. In addition, since the viscosity increases as crosslinking (stabilization) proceeds, the mobility and growth rate of the LMOs decrease; hence their final size remains small. The smaller the size of the LMOs, the
la
c
c
c
c
N
N
N
N
Ill
H
Ill
Ill
H
H
H
H
I
I
I
I
I
l
l
Ill
H
H
H /:\> % ;\/;\;/:
H
'LL,
1
/c\N/c\N/c\N/c\N
+ + + +
0
0
179
0
/
IV
0
Fig. 9.8 Incorporation of oxygen into cyclized PAN fiber (Clarke and Bailey, 1973).
180 Cavbonfibers less graphitizable is the carbon and the lower the properties of the fiber. In order to ensure appropriate N/C ratios at reasonable temperatures, stabilization should result in only a moderate degree of crosslinking. In addition, slow heating rates during precarbonization should permit hydrogen and delay nitrogen emissions; both of these effects lower the temperature at which extensive formation and growth of LMOs occur. A commercially acceptable rate of stabilization requires the use of as high a temperature as possible. However, since the reactions that occur during stabilization are exothermic, it is most important to limit the oxidation rate and to prevent uncontrollable temperature increases. These conflicting requirements have resulted in the development of alternative methods of stabilization. These include stabilization in: hydroxylamine solution (US Patent 3 767 773, 1973), aminophenoquinones (US Patent 4 004 053, 1976), aminosiloxanes (US Patent 4 009 248,1977), amine salts (US Patent 4 009 248, 1977; US Patent 4 024 227, 1978; US Patent 4 031 288, 1978) or stabilization in gas phases such as mixtures of NO and NO,, Br, and 0,, or HC1 and 0,. Other stabilization processes have been proposed that are designed to reduce the cost and/or decrease the stabilization time. It seems possible to reduce the time by stabilizing the fibers in persulphate (US Patent 3 650 668, 1972), cobalt salt (US Patent 3 656 882,1972), nitric acid (US Patent 3814377, 1974; US Patent 3 656 883, 1972), or to control the final quality of the fiber by stabilizing in carboxylic acid (US Patent 3 814 377, 1974; US Patent 3 656 883,1972), or nitrophenol. Processes designed to produce fibers of larger modulus have also been developed. These involve stretching the precursor during stabilization (US Patent 3 917 776,1975; US Patent 3 677 705,1976). Fiber manufacturers attempt to fit the physico-chemical conditions of the various operations cited above (nature and proportion of the co-monomers, cyclization, stretching, stabilization, carbonization) to obtain an opti-
mal product, i.e. forming the LMOs at the smallest reasonable temperature, retaining the largest nitrogen content beyond the temperature of LMO formation and incorporating an optimum amount of oxygen during stabilization to prevent polymer melting without inducing too small a LMO size. It is known that the overall oxygen content should be between 8 and 12 wt.% in order to completely stabilize PAN fibers (US Patent 4069 297, 1978). Less than 8 wt.% oxygen gives a large weight loss on carbonization due to excessive evolution of gases from the incompletely stabilized central core; more than 12 wt.% oxygen degrades surface layers and the properties of the final fiber (Johnson, Rose and Scott, 1970). Exactly what an average value of 8 wt.% translates into for the specific oxygen content of the surface layers and the core region is unknown, but it would obviously depend on fiber diameter and the kinetics of the stabilization process (diffusion or reaction controlled). A large diameter PAN fiber containing an average of 8 wt.% oxygen exhibiting diffusion controlled stabilization kinetics would probably be composed of highly degraded surface layers with perhaps an under-stabilized central core. Conversely, a very thin fiber exhibiting reaction controlled stabilization kinetics might be completely and homogeneously stabilized with an average oxygen content of less than 8wt.%. Presumably a similar statement can be made for the stabilization of pitch fibers. In the stabilization process, the effect of fiber diameter on the rate of oxygen uptake is important. The three curves shown in Fig. 9.9 illustrate the slower rate of oxygen uptake exhibited by fatter fibers. In addition, larger diameter pitch fibers tend to exhibit a diffusion controlled stabilization that produces an under-stabilized central core and an over-stabilized fiber skin. Smaller diameter pitch fibers appear to exhibit reaction controlled stabilization since no skincore type microstructures are observed.
Chemical changes during carbonization 181
....................
________---------
Solid line - mesophasic particles
-4.881
-5.08 8.68
( O x i d i z e d Oxygen, "
"
I
'
"
'
I
"
"
388 Deg I
'
Time (hours)
Fig. 9.9 Effect of diameter on the stabilization of mesophasic particles and pitch-based fibers (Kowbel, Wapner, Wright, 1989). 9.6 CHEMICAL CHANGES DURING CARBONIZATION
The carbonization of stabilized PAN and pitch involves controlled heating to a temperature of about 1500°C.The majority of gases emitted from either the PAN or the pitch are emitted before a temperature of 1000°C is reached and the emission is primarily from unstabilized regions (Jain and Alhiraman, 1987; Lewis, 1982). Indeed, the quantity of gases emitted from an unstabilized central core of either PAN or pitch can be so large that the fiber can disintegrate. Great care should therefore be taken to determine the optimum heating rate for stabilized or under-stabilized fibers. In some cases, hold times should be incorporated into the heating cycle. Both materials emit a variety of gas molecules containing oxygen, hydrogen and carbon; however, a major difference between PAN and pitch involves nitrogen containing compounds which are only emitted from PAN. The temperature and
rate of emittance are important control parameters since they affect the strength of the resultant carbon fibers. A stabilized polyacrylonitrile fiber which contains about 11wt.% oxygen can be thermally degraded by heating at a slow heating rate (Riggs, Shuford and Lewis, 1982) in an inert atmosphere such as nitrogen or a reactive environment where nitrogen gas is bubbled through acid (US Patent 3 972 984, 1976) or water (US Patent 3 677 705,1976; US 3 656 903, 1972; US Patent 4 039 341, 1976).As the temperature increases, many complex reactions take place resulting in the evolution of volatile products. For example, when the fiber is initially heated, cyclization occurs with the release of large amounts of HCN and NH,. Up to 450°C, HCN, acrylonitrile, propionitrile, NH, and H,O are emitted. Subsequently, at around 500°C and 700°C copious quantities of HCN and water vapor are emitted, respectively. All of these emissions are believed to
182 Carbon fibers come from reactions involving crosslinking of individual molecules. Evolution of nitrogen starts near 700°C; so fibers produced after being heated to 1000°C retain only about 5.8 wt.% nitrogen and have lost about 50 wt.% of the original PAN precursor fiber. Results obtained from experiments involving slow pyrolysis at 4"C/min indicate that optimum mechanical and physical properties are unobtainable unless high nitrogen contents are retained within the precursor until the later stages of carbonization (Deurberque, 1990; Deurberque and Oberlin, 1991). Therefore, a large nitrogen content (large N/C ratio) should be present when local molecular ordering (LMO)begins and the carbon skeleton is being formed. Since the N/C atomic ratio depends inversely upon the H/C ratio, LMO should occur at large N/C and small H/C ratios. Essentially, small amounts of aromatic hydrogen and a relatively large amount of nitrogen present during the LMO stage allow the carbon skeleton to remain flexible enough at high temperatures that molecular rearrangement is easy. Within this overall fibrous texture, the nanotextural features of the carbonized fibers are the consequence of the variations in cyclization, stabilization, carbonization and graphitization conditions. If the original precursor is CH- and NH-rich but oxygen-poor, the corresponding carbonized fibers will have low porosity, high compactibility and stacking order and a relatively high strength. Likewise, if two nitrogen atoms are present in two aromatic rings contained within adjacent sheets, they are able to promote bonding of contacting BSUs together with a N, release (Watt, 1972). The ultimate strength value of the fibers increases as the compactibility and the availability of 'efficient' nitrogen (i.e. the nitrogen remaining at the moment of LMO occurrence) increases (Oberlin and Guigon, 1988; Guigon, 1985). Bright and Singer (1979) agreed with others when they found that the tensile strength of most heat treated fibers tends to decrease with higher temperature exposures and release of nitrogen (Fig. 9.10). The same
40 .1-
3.5 -
100
3.0-
.-
c
0-
too
z
G2
2.5-
W
W
E
a
t;; W
z w
n
I
I
f
= I)
g (3
2.0-
300
W
$ 2 W
t-
k
1.5 -
1.0
zoo
-
Fig. 9.10 Effect of heat treatment temperature on the tensile strength of carbon fibers (Bright and Singer, 1979).
authors argue that if adequate nitrogen exists within the fiber after LMO occurs, then a greater potential for bonding exists and an improved rather than a diminished strength will result. This conclusion is based on the observation that, in contrast to other fibers, the strength of (nitrogen containing) commercially available Toray T-300 fibers increased from 2.2 to 3.2GPa after heat treatment to 2800°C. Tension also increases cyclization and nitrogen elimination which increases the tensile strength of the final carbon fiber (Watt, 1972). Exactly what influence nitrogen has on the development of high tensile strength is still being debated. However, the texture of the graphitic layers, the amount of nitrogen present in the original precursor and the temperature at which misshapen layers touch -
Microstructural changes during carbonization 183 and hence are available to bond and emit nitrogen - are obviously important. Up to lOOO"C, carbonization leads to an effluentloss and increasing aromatization. As a result, the solid residue transforms from a viscoelastic into a brittle solid material. The stabilized PAN normally carbonizes into a statistically isotropic but nanoporous carbon material which, because of the small dimensions of the initial LMOs, is inherently non-graphitizable (Joseph and Oberlin, 1983a). Continuing pyrolysis up to 1500°C eliminates most of the residual nitrogen and completes the conversion of the PAN molecules into sheets of carbon that are appreciably anisotropic (Mair and Mansfield, 1987).Continued heating eliminates the remaining nitrogen and, since the material is then only made of pure carbon, further modifications are only structural. Graphitization is the name of the process that involves heating the carbonized fiber to approximately 2500°C in times as short as a minute (US Patent 4 005 183,1977).Graphitized pitch fibers exhibit a larger, more graphitic and better oriented crystal structure than PAN-based carbon fibers which are inherently non-graphitizable. Parallel to the fiber axes, pitch fibers have higher stiffness and thermal conductivity values and a reduced thermal expansion coefficient. These changes due to graphitization do not produce any significant increase in relative strength values. As a result of the extreme temperatures required to process them, graphitized pitch-based carbon fibers are more expensive and are fabricated for specialized applications.
Fig. 9.11 Model of crumpled sheet-like structure (Guigon,Oberlin and Desarmot, 1984a).
temperatures, for the bonding of adjacent sheets. These sheets contain numerous vacancy imperfections and are folded to enclose pencil-shaped voids oriented in the general direction of the fiber axis. The lengths of each block or sheet are relatively short, with each 9.7 MICROSTRUCTURAL CHANGES DURING succeedingblock misoriented with respect to its CARBONIZATION neighbor. A schematic illustration of the The initial heating of stabilized PAN fibers microstructure as it exists within the actual causes growth of graphite-like ribbons by a fiber is shown in Fig. 9.12. This structure is typdehydrogenation mechanism. Denitrogenation, ically exhibited by high strength PAN-based which occurs as the temperature is increased, is carbon fibers. Further temperature increases tend to responsible for the growth in area and the transformation of these ribbons into thin sheet- decrease the void space by joining sequentially like structures (Fig. 9.11) and, at higher oriented and touching graphite like layers and
184 Carbon fibers
fibre axis
Fig. 9.12 Model of microstructure within a high strength PAN-based carbon fiber (Guigon, Oberlin
and Desarmot, 1984a). aligning them more parallel to the fiber axis. The distorted sheets of BSUs associated with tilt and twist boundaries are bonded to each other wherever the boundaries of adjacent sheets touch. The more compact the fiber, the larger the number of contact areas and the Fig. 9.13 Model of PAN-based high modulus cargreater the chance for adjacent sheets to bond. bon fiber (Guigon,Oberlin and Desarmot, 1984b). The lateral cohesion thus formed causes the strength of the fiber to increase. Specifically, the distortions within the polyaromatic centrations and, hence, weaker fibers graphene layers or sheets tend to be removed (Reynolds and Sharp, 1974; Reynolds and by accumulating any structural defects at their Moreton, 1980). boundaries. This induces a progressive increase of the width and the radius of curva9.8 ELECTRICAL AND THERMAL ture of the aromatic layers which can be PROPERTIES correlated with the stiffness, stacking order and the diameter of the graphitic layers Studies of microstructural features have been (Oberlin, 1984).A schematic of this microstruc- carried out using techniques that include: ture is shown in Fig. 9.13. The longer, better X-ray and electron diffraction, electron spin oriented and more graphitic microstructures resonance, thermoelectric power, magneto exhibit both higher values of moduIus and resistance, lattice fringe imaging, etc. All thermal and electrical conductivities; unfortu- results indicated that carbon fibers are comnately, the misalignment of the larger posed of turbostratic layers of graphite microstructural units causes large stress con- oriented preferentially at some angle to the
Electrical and thermal properties 185 fiber axis. Increasing the heat treatment temperature results in a reduction of the interlayer spacing, a decrease in void space, a growth in thickness and area of the graphitic crystallites and an increase in the preferred orientation of the microstructure. All of these changes increase the elastic modulus and the electrical and thermal conductance. A corresponding reduction of the tensile strength also occurs by mechanisms that depend on local defects as discussed in the previous section. A comparison of the g-value anisotropy of pitch and PAN in Fig. 9.14 indicates that the degree of anisotropy changes for both fibers after heating to about 1700°C. Although pitch based fibers become more anisotropic when the temperature is increased further, the anisotropicity of PAN seems to saturate at a level which is comparable to that of a pitch fiber heated only to about 2000°C. A simple consequence of this inability to fully graphitize
x104
140
-
t
Long Heat Treatment
:120100-
"
\
I
/
1
I
/ I /
0)
dl
PAN-based fibers is the lower maximum values of the elastic modulus and the electrical and thermal conductivities. Typical results which compare the effect of heat treatment on the electrical and thermal conductivities are shown in Figs. 9.15 and 9.16.
/
1800 2200 2600 3000 Heat Treatment Temperature, deg C
Fig. 9.14 Variation in g-value anisotropy of pitch-base and PAN-base carbon fibers as a function of heat-treatment temperature (HTT) (Aggarival, 1977).
Single Crystal Graphite ----------------
1000
1500
2000
2500
3000
35 0
HEAT TREATMENT TEMPERATURE (OC)
Fig. 9.15 Schematic variation of the room temperature electrical resistivity against T,, for: (0) ex-rayon; ( 0 )hot stretched rayon; (A and V) ex-PAN fibers (Robson et al. 1972); (0) Ex-pitch fibers (Bright and Singer, 1979)and (solid curve) benzene-derived fibers (Chieu et al. 1983). The scatter of typical data points about the mean give an indication of the uncertainty. The dashed line indicates the decrease in resistivity produced by hot stretching the ex-rayon fibers.
186 Carbon fibers E. GPa
l
tiI '
0.01 1
'
20
I
I
30
40
I u)
Ex
I
I
I
I
60
70
80
90
1
106 psi
Fig. 9.16 Conductivity of pitch-base, PAN-base and rayon-base carbon fibers as a function of the tensile modulus of elasticity. (Courtesyof R. Gray, NSWC, Dahlgren, Virginia.)
9.9 MECHANICAL PROPERTIES OF FIBERS
The microstructural changes discussed above have been deduced using X-ray diffraction techniques. In addition, the mean length of the graphite sheets oriented in the fiber direction La and their thickness Lc may be computed using such techniques. It has been found that both of these parameters increase with increasing temperature. The orientation of these layers also becomes increasingly aligned with the fiber axis. The net effect is to increase the tensile modulus continuously as shown in Fig. 9.17. Conversely, the tensile strength tends to decrease (Fig. 9.10). More recent studies, which allow direct
observation, have used electron diffraction (Guigon, 1985) and lattice fringe imaging (Oberlin and Guigon, 1988). This latter technique is particularly powerful, because in dark field, the parallel orientations (with respect to the imaging beam) of the convoluted graphite layers can be imaged. Sketches of the possible layer convolutions, the images they can produce and the measurements that can be made are shown in Fig. 9.18. Most important is the realization that it may be possible to classify the properties of commercial samples in terms of decreasing either transverse (rt) or longitudinal (r,) radius of curvature of the sheets. Such a classification, based on these and other numerical values
Mechanical properfies offibers 187
I
1000
2000
3000
shown in Table 9.2, could form the basis of empirical relationships between micro-texture and mechanical properties. At increasingly higher heat treatment temperatures, the scattering domains within high modulus fibers become large and well defined so that the length of the graphitic sheets in the fiber direction, .La,,Cr can be measured directly (see Table 9.2) from the observed Moire fringes. Correspondingly, the radii r, and rI of the sheets are also measurable from 002 lattice fringes. The lateral cohesion of the fiber is also ensured by bonding between adjacent distorted sheets of carbon wherever two grain boundaries are in contact. The chances of such bonding increase as r, and rI decrease, but decrease as La increases. Hence the extent of lateral cohesion can be defined by a variable S = La [(l/rf) (l/rt + l/rl)]. As shown in Fig. 9.19, a linear correlation has been observed between CY, and 1/ S indicating that long, relatively unbonded graphitic layers result in weak fibers. Conversely, Young's modulus and electrical
1 3500
THT ( ' C )
Fig. 9.17 Effect of heat treatment temperature on the elastic modulus of PAN and pitch-based carbon fibers. (Data from Johnson, 1969 and Aggarival, 1977). 1
m 2rr sin 37.
3
2
n
m
n
2r,sin37*
Fig. 9.18 Sketches of the possible dark-field images of a longitudinal section (lamellar model of Fig. 9.15) as r decreases and the fold develops (Guigon, Oberlin and Desarmot, 1984b).
188 Carbon fibers Table 9.2 Quantitative measurements which are suggested for use to classify microtexture and mechanical properties SAD patterns 002DF ~ _ _
[O]
La
__
N L,
Number of fringes in a stack Length of a perfect fringe Length of a distorted fringe
L?
__
11DF
__
Thickness of elementary bright domain [BSU] Length of elementary bright domain
Lc
002LF __
Half width of 004 reflections Half width of 11[0] ring
L'coo4
L'd,
Length of turbostratic Moire fringes Diameter of a domain showing rotational Moire fringes
Lall TF Lall Cr
~-
P
ODP of 002LF
Arc opening Interfringe spacing spreading
ADO,,
,,
lt +
0
I
1
1
I
"
'
I
I
I
1
0.1
1
'
f
i
I
I
'
I
I
I 0.2
11s
AA-
Fig. 9.19 Numerical relations between tensile strength and the microtexture oc = f(l/S). High modulus fibers (full line). High tensile strength fibers heat treated at 2800°C (dashed line) (Oberlin and Guigon, 1988).
conductivity correlate well with La, as seen in Fig. 9.20. 9.9.1 MICROSTRUCTURAL CONSIDERATIONS
According to the pioneering work of Griffith (1920), the following expression describes the strength, oF,of a brittle solid, containing a crack of length 2 4
oF=d(Ey/4a) (9.1) Carbon is a brittle solid; thus, since no plastic deformation can occur, very high local stresses where E is the elastic modulus and y is the surwill develop at stress raisers, such as disconti- face energy. Inspection of equation (9.1) nuities, changes of section size, cracks, etc. indicates that longer cracks are more effective
Mechanical properties offibers
189
(a)
I
I
longer cracks generate higher tip stresses, it can be inferred that once a crack begins to move it will continue to move (accelerate) until it reaches the geometric boundaries of the material. As a consequence of this, the failure of brittle solids is abrupt and depends on the probability that a crack of some critical length is present. If such a theory can be applied to carbon fibers, then it can be argued that the maximum length of cracks (or similar microscopic stress raisers) that can be contained is limited by the fiber diameter; so small diameter fibers will be stronger than large diameter fibers. In addition, since the chance of a crack being present is greater, long fibers will tend to be weaker than short fibers.
I
I
iil
In a composite, many fibers are arranged more or less parallel to one another and function as a load bearing bundle. A number of publications have appeared (Herring, 1966; Wright and Iannuzzi, 1973; Wright and Wills, 1974) that discuss the distribution of strengths exhibited by brittle fibers and how these distributions can be used to compute a mean strength, om and a corresponding bundle strength, ob. For example, the above authors argued that their individual fiber strength data tended to obey a Weibull distribution characterized by the expression, G(o) = 1- exp{- a(o/oo)")
(9.2)
where G(o) is the probability of failure of a fiber subjected to stress o,oois the distribution
190 Carbon fibers scale factor, w is the distribution shape factor and a is a function of the length/diameter (L/d) ratio of the fiber (Corten, 1967). If many fibers are tested of different length I, then a graphical method can be used to deduce w, a and a. and, the mean strength and the strength of a bundle of fibers can be computed from
om=oo(i/d)-lw(i + i/c~),
(9.3)
where r is the gamma function and Gb=oo(acl)e)-'lY
(9.4)
Bundles of twisted fibers would exhibit lower strengths. 9.10 COMPOSITES FABRICATED FROM
CARBON FIBERS
tion of the same fiber by a shear process. The length of matrix required to do this defines a bundle of short fibers (or segment of composite) which must break in order to break the composite. Such bundles can be modeled as analogous to a link within a chain; failure of the weakest link defines the failure load of the composite. Nevertheless, since shorter fiber bundles are stronger than longer, shorter links are stronger than longer links. The link length, 6, has been discussed by Rosen (1964) and, for purposes of this discussion can be approximated by,
6=~ d / 2 ~
(9.5)
where o = obthe stress in the fibers at failure of the composite, d is the diameter of the fiber and z is the shear strength of the matrix or matrix-fiber bond as it exists in the composite. For a given fiber strength distribution, the stronger composites will all exhibit smaller ineffective (link) lengths. This is accomplished by using small diameter fibers, well bonded using high strength glue.
Carbon fibers are very strong, stiff and lightweight materials. In addition, their small diameter (8-12 p)makes them extremely flexible. Unfortunately, they exhibit little compressive strength and they exhibit a poor abrasion resistance. A solution to these problems, and to the problem of brittleness, is to 9.10.1 SURFACE TREATMENT bond large numbers of fibers together to form a composite solid. In this case, the glue or A freshly prepared fiber does not bond well to bonding agent forms a continuous phase that a polymeric glue (or to anything else for that is usually defined as the matrix. The matter); however, the tendency to bond can be matrix-fiber mixture is called a composite significantly increased by subjecting the fiber material. The function of the matrix is to sup- surface to a controlled oxidation. As discussed port and separate the fibers, to protect them by Eggs, Shuford and Lewis (1982) in the earfrom reaction with the environment and to lier edition of this handbook, this treatment transfer load. In a composite, the tensile and essentially etches the surface, cleans it, compression properties parallel to the fibers increases its surface area and produces polar are much better than those measured on bun- hydrophilic oxygen-containing groups which dles. The transverse properties are also bond to it. The process can be carried out in a optimized, since the matrix serves to improve liquid or gaseous environment; for example, the fiber-matrix connectivity. This function is heating in air or oxygen-nitrogen mixtures, important since it affects the mechanical and CO,, C1, NOz-NO, NH, and plasma-ionized thermal properties in the transverse direction. inert gases oohnson 1969; US Patent 3 754 957, Fortunately, although the strength of individ- 1973a; US Patent 3 723 150, 1973b; British ual fibers exhibits a pronounced size effect, no Patent 1341 161, 1973; U.S Patent 4 374 114, size effect is exhibited by composites. In effect, 1983; US Patent 3 627 466, 1971; US Patent load is transferred around fiber breaks into 3767774, 1973; US Patent 3780255, 1973). adjacent fibers and back into the unbroken sec- Direct wet chemical oxidation has been tried
Mechanical properfies of unidirectional composites 191 using aqueous nitric acid, hypochlorite, chlorate and dichromate in sulfuric acid. Treatments have also been investigated using electrolytes of hypochlorite, ammonium hydroxide, sodium hydroxide and ammonium sulfate (US Patent 3 660 140, 1972; US Patent 3 746 506,1973; US Patent 3 894 884,1975a; US Patent 3 859 187, 197513; US Patent 3 746 450, 1973; US Patent 3989802, 1976; US Patent 3832297, 1974; US Patent 3671422, 1972; British Patent 1371621, 1974; British Patent 2 071 702,1981) . During oxidation, the strong carbon-oxygen complexes which are formed bond tenaciously to the fiber surface and will subsequently react with a matrix resin. In order to preserve this reactivity a thin layer of the final matrix resin is applied to the surface of the fibers as a finish or size. This layer does double duty in both protecting the fiber surface against damage during transportation, further processing and handling and in promoting wetting when the sized fibers are bonded together with the matrix resin. Poor bonding is a sensitive function of the surface morphology, anisotropicity, heterogeneity and the nature of the interphase layer between the fiber and the matrix. For example, it has been found that the greater the degree of graphitization and the better the alignment of the microstructure with respect to the fiber axis, the poorer the fiber will bond. Essentially, the higher modulus carbon fibers will not bond easily in the absence of a surface treatment. In some composites where the failure strain of the matrix is smaller than the failure strain of the fibers (as it is for ceramics or carbon), poor bonding is an asset since the largest fiber-matrix bond strength is not required. Conversely, if the failure strain of the matrix is larger than that of the fibers, a strong bond is desired. The reason for this apparent dichotomy involves the fact that fibers are added to brittle matrices primarily to toughen them; only tough matrices (large failure strain) can be strengthened. The mechanism of toughening depends on the blunting of a running
crack by interaction with low strength fiber-matrix bonds; strengthening depends on load being transferred from the matrix to the fibers through a strong fiber-matrix bond.
9.11 MECHANICAL PROPERTIES OF UNIDIRECTIONALCOMPOSITES
The properties of unidirectionally reinforced composites are strongly orthotropic. Specifically, properties measured parallel to the fibers are quite different from those measured at right angles to them. More importantly perhaps is the sensitivity of property determination with respect to the direction of the measurement. For example, Fig. 9.21 illustrates the change in elastic properties exhibited when a unidirectionally reinforced composite is loaded at some angle to the fiber axis. In this case, it can be observed that the stiffnesses of the composite (Qll,etc.) begin to change significantly when the load is misaligned only fifteen degrees to the fiber axis.
9.11.1 MICROMECHANICS
Using micromechanics, various equations can be used to estimate the properties that might be exhibited by well bonded composites. For example, parallel to the fibers, the modulus (E,), strength (0,) and Poisson's ratio y2are given by;
E , =Vf E, + Vm Em
(9.6)
ol=vfof+ vm0,
(9.7)
'12
= Vfv12
+
'mum
o2z om
(9.8) (9.9)
where V is the volume fraction of fibers (0 or matrix (m) respectively, ofis the strength of a bundle of fibers with length equal to the ineffective length (usually, due to lack of statistical data ofis taken as the mean strength supplied by the fiber processor). Also,
192 Carbonfibers
9.11.2 MACROMECHANICS OF LAMINAE
There are a number of excellent books and monographs that have been written on macro-
I
30
01
[
60
1
60
90
-8
mechanics of composites; thus, the reader is directed to these for a more complete discussion of this subject (Ashton, Halpin and Petit, 1969; Jones, 1975; Tsai and Hahn, 1980; Daniel and Ishai, 1994). When only plane stress conditions exist, (e.g. 03=0 and E, is related to E, and E,), then it is possible to relate stress to strain along the principal axes of an orthotropic lamina,i.e. parallel (1)and perpendicular (2) to the fiber axis:
I
01
30
60
'
90
-6
(42
I
01
I
a
30
60
90
-
8
Fig. 9.21 Transformed, off-axis modulus of T300/5208. The angle is the ply orientation and is positive for counterclockwise rotation (Tsai and Hahn, 1980).
Mechanical properties of unidirectional composites 193
(9.12)
the components of the [Q],, has already been illustrated in Fig. 9.21. (9.15)
or
where
and
(9.13) In addition to the variation of elastic properties, the strength of unidirectionally reinforced composites has been found to be sensitively dependent on the angle of loading. A typical failure curve for tension and compression is shown in Fig. 9.22. This type of curve can be obtained by transforming the applied stresses to directions parallel and perpendicular to the fiber axes and then equating those stresses to the failure strengths actually measured (or computed using micromechanics) along those directions. Failure of a composite is then considered to occur when the transformed stress exceeds the failure stress actually measured in that direction, i.e. when
Inspection of equations (9.12) and (9.13) indicates that parallel and perpendicular to the fiber axes, tensile (compressive)stress produces tensile (compressive) strain and shear stress produces shear strain. There is no coupling between shear stress and tensile (compressive) strain, e.g. Q,, = 0 = Q,, as shown previously in Fig. 9.21 and the shear coupling coefficient which is the ratio of tensile (compressive)stress to shear strain, is zero. The above conclusion is not valid when loads are applied at some angle to the principal axes since appreciable shear coupling can occur; Q,, and Q,, are not zero. This means that tensile (compressive) stresses produce shear strains in addition to the more normal tensile (compressive strains). The stress-strain relationships are then written, (9.14)
where x and y are the orthogonal axes of the composite test specimen that are oriented at some angle to the fiber axes and where [Q] or .xY [SI, are the transformed matrixes of equabons (9.12) and (9.13). The variation with angle of
F,,, = Fl,/(cos28);Fx,= F,, / (sin20); F,, = F6/ (sine c o d ) and
Fxtc = FlC/(~os20); Fxc = F J ( s i n 0 ~ 0 s 0 ) and F, are the measured failure where Fl,(c,,F2t(c) tensile (t) or compressive (c) loads measured parallel (1)or perpendicular (2) to the fiber axis and F, is the measured shear strength. These failure criteria are collectively described as the maximum stress failure criterion. There are other failure criteria such as maximum strain, deviatorial strain energy (Tsai-Hill) and Interactive Tensor Polynomial (Tsai-Wu) that can be used, some of which allow for interaction between the stresses. However, for typical composite structures, the maximum stress criterion give reasonably conservative estimates of expected failure stresses. 9.11.3 MACROMECHANICS OF LAMINATES
In order to eliminate coupling and to reduce the very strong change of mechanical properties
194 Carbonfibers 1.2 1.o a 0.8 a (3 0.6 LL'
160 120
.-
(I)
80 x-
0.4
LX
40
g-
F! = o
O
E
5 -0.2
-40
g
0.2
-0.4 -0.6
m
i5
-80
-0.8
0
10
40 50 60 70 Fiber orientation, 8, deg.
20 30
80
90
Fig. 9.22 Uniaxial strength of off-axis E-glass/epoxy unidirectional lamina as a function of fiber orientation (Daniel and Ishai, 1994).
with the direction of applied load, composites are usually fabricated from multiple layers, each arranged at some angle to their neighbor. The angles that are used can be arranged to bring the fiber axes at some optimum angle to the expected loads; however, there is a requirement for each lamina to be oriented very precisely with respect to one another in order to avoid tensile-shear effects and coupling between in-plane loading and out-of-plane deformation (tensile loads can be made to produce bending and twisting deformation, for example). Various rules of angle-ply materials have been worked out in order to avoid the cross-coupling terms. For example, symmetric laminates exhibit no coupling between in-plane loading and out-of-plane deformation. (Symmetric laminates define a composite in which for every lamina oriented at some angle there is another layer of identical thickness and orientation placed at an equal distance from the mid-plane of the composite.) And, composites which exhibit no shear coupling are 'balanced laminates' (pairs of identical layers oriented at a positive and an equal negative angle with respect to the laminate reference axes).
Of particular interest is the quasi-isotropic composite which is often specified for commercial structures. The in-plane engineering elastic constants of these materials are identical in all directions and there is no shear coupling. Examples of such composites are symmetric arrangements of [0/60/-601 or [0/ +45/90] layers. 9.12 TESTING TECHNIQUES
A knowledge of the behavior of the constituent phases allows the mechanical properties of the resulting composites to be computed. These calculations are made with the use of expressions (9.6)to (9.11).The results of the following tests should provide input to and in some cases confirm these calculations. The following was adapted from the excellent (and much more extensive) discussion contained in Chapter 8 of the book Engineering Mechanics of Composite Materials (Daniel and 0.Ishai, 1994). 9.12.1 PROPERTIES OF FIBERS AND MATRICES
Determination of the elastic and failure properties of the fibers is described in ASTM
Testing techniques 195 specification D3379. The difficulty is mostly verse Young’s modulus, tensile strengths and involved in determining the elastic displace- strains and the major and minor Poisson’s ment parallel to the fiber axis, since no ratio can all be determined from this type of measurement device can be attached directly coupon test specimen. Similar properties can to the fragile fiber. Since the fiber tends to shat- also be measured in compression; usually ter and disintegrate when failure occurs, the however, short, thick specimens are used in mean diameter measured before the test has to order to avoid buckling failure. ASTM D-3410 be used to calculate the failure stress. describes this test method. The method involves attaching the fiber across a slot cut into a paper tab. The composFlexure testing ite specimen is aligned coincident with the load axis of the testing machine, the tab ends A far more expensive test is described in ASTM are gripped and the sides of the tab are cut to C393. This requires a rather large allow only the fiber to be loaded by the (22 in/l in/1.5 in) sandwich flexure specimen machine. The measured compliance Cm is the which is tested in four point bending. The sum of the compliance of the loading system dimension of the honeycomb core and the composite face sheets are adjusted to cause failC, and the compliance of the fiber C,. Thus ure in the approximate face sheet. Good results Cm= C,+ C, = C, + (AE,) can be obtained in both tension and compreswhere 1 is the fiber length and E , is the fiber sion; however, since failure might occur in the modulus. A plot of measured compliance core or in either of the outer skins or at the against fiber length allows calculation of the skin/core interface, care should be exercised in loading system compliance and the fiber both determining the exact failure mode and in reporting the appropriate failure stresses. modulus. In the above paragraphs the reader has Polymer matrices are evaluated using coupons cut from thin sheets. Typical geome- been cautioned about making sure that the tries are described in ASTM specifications loading axis of the testing machine and the symmetry axis of the specimen are coincident. D638, D638 and D882. Indeed, consistent measurements will only be achieved by eliminating any tendency to develop complex stresses in any region of the 9.12.2 PROPERTIES OF COMPOSITES specimen. An additional effect not significant Coupon tests in designing or testing ductile materials is the The determination of the tensile longitudinal tendency for composites to exhibit poor shear and transverse properties of unidirectionally properties. Indeed it is common for engineers reinforced composites can be obtained by test- designing metallic components to assume that ing relatively long coupon specimens. These failure in shear will not occur if failure in tenspecimens, as described in ASTM specifica- sion or compression is designed against. tions D3039-76, are 9 in long, 0.5 in wide and Unfortunately, shear resistance of composites from 0.02 to 0.10 in thick. The apparently is not directly related to tensile or compressive excessive length requirement is an attempt to properties; thus, a separate shear failure criteminimize the effect of specimen misalignment rion must be used when designing with with respect to the loading axis. Glass/epoxy composites. Shear failure can occur in-plane or tabs are bonded to the specimen ends to shear interlamina. the load into the specimen and to avoid damage and failure of the specimen within the gripped length. The longitudinal and trans-
,
196 Carbon fibers
In-plane shear testing A reasonably simple test method which utilizes a coupon with an eight layer symmetric +. 45 deg. layup is described in ASTM standard D35 18. If the specimen has strain gages oriented parallel to the x and y axes of the specimen (not the fibers), then the failure stress in shear is given by (0JmaX/2, the shear failure strain is ( E ~- &Jmax and the shear modulus is given by GI,= E J ~ ( E-~ E,) A rail shear test has been described in ASTM D4255-83. The shear stress and shear strain can be obtained at intermediate and maximum loads. Thus, GI, and the appropriate failure parameters can be obtained.
Interlaminar shear strength The interlaminar shear strength is a measure of the strength of the bond that exists between the various layers within the composite. It is important to know such a value since bending of beams can cause appreciable shear stresses, which while not large enough to cause failure of a traditional metallic structural material can fail a composite. This occurs either within the matrix or at the fiber-matrix bond line. The test specimen, as discussed in ASTM D2344, is a rather short, thick beam which is tested in three-point bending. Basically, the test specimen is sized such that the ratio of the shear stress generated at the midplane is maximized with respect to the tensile or compressive stresses generated in the outer fibers. If inspection of the failed specimen indicates that failure by shear has occurred, then beam theory indicates that a reasonably accurate estimate of the interlaminar shear strength F,, can be calculated from F3,=3P/4wh where P is the maximum load and w and h are the width and depth of the beam. Another way to measure the same maximum shear stress is by using the double-notched specimen as described ASTM
D3846. This specimen contains two square notches, each cut into an opposite face of a tensile or compressive specimen. If the slots are cut a reasonable distance, I, apart in order to avoid interactions between the stress fields of the cut slots, then the interlaminar shear strength is given by
F,, = P/wl
9.13 CONCLUSION AND PERSPECTIVES
Low performance isotropic non-continuous pitch-based fibers and anisotropic PAN-based large tow fibers have been available for a number of years in the USA priced at $20/kg. These fibers have been used to produce electrically conducting flexible heating materials to control the temperature of pipelines, to heat human dwellings, hot-houses, etc. (Karpinos and Izmalkov, 1982; Glushchenko and Griffen, 1982; Levit, 1986; Jakubowski and Subramanian, 1979).In addition, low modulus carbon fiber materials and their composites have been used as abrasive, anti-friction, sealing and heat-insulating materials. Other nonstructural uses include activated carbon fibers which can be used directly for the purlfication of nitrogen oxide-containing fumes, chimney smoke, automobile air conditioners, various respiratory, water purification (Chupolov et al. 1983; Richter, Knoblauch and Juntgen, 1984). In recent years carbon fiber adsorbents have been used in medicine to remove toxic substances from body fluids (Ternovoiet al., 1985). The high modulus/high strength continuous carbon fibers have been available at a cost of $60-80/kg. Thus, reinforced epoxies, polyesters and other polymers have demonstrated properties that have enabled them to find use in the aerospace and aeronautical fields. Smaller quantities of fibers have been used in high quality sports gear such as golf clubs, fishing rods, tennis rackets, marine sports, ski equipment, bicycle equipment, etc. Fibers have also been used in many civil engineering
Conclusion structures. These include short fiber reinforcement of cement mortar and continuous reinforcementof concrete. Some attempts have been used to provide earthquake resistant structures and fiber reinforced ropes and some successes have been reported in fabricating structures with carbon fiber reinforced aluminum and copper. Carbon fiber reinforced carbon has been considered for high temperature load bearing structures and heat shields in spacecraft and supersonic aircraft. And, finally, similar materials are considered useful in nuclear applications and disc brake materials for aircraft, high speed trains and racing cars. At the present time, there is a growing interest in the use of very high thermal conduction properties to manage local temperatures in sensitive electronic equipment. Nevertheless, despite all these apparent successes, it is vitally important to realize that the expanded use of high performance carbon fibers depends very sensitively on the lifetime costs involved in substituting carbon and its composites for competing metals, ceramics and polymers. It has been recognized for many years that market penetration of carbon fibers will always be limited to rather sophisticated structures if the cost remains at the present high level. However, the precursor PAN material presently costs about $5/kg; thus, taking into account the weight loss and processing costs involved in converting PAN to carbon, it is unlikely that large amounts of carbon fiber made from such material will ever be less than $10/kg. Conversely, the cost of the pitch precursor material is almost insignificant, since it is the byproduct of a commercial process established to produce other end products: gasoline, metallurgical coke, etc. In addition, since the fiber is fabricated using a melt spinning process, the production rate can be much faster than the wet or dry process used to produce PAN based fibers. The carbon yield from pitch precursors can average up to 85%, whereas the carbon yield produced from PAN averages about 65%. Pitch based carbon fibers
197
can be made with modulus values much larger than can be obtained from PAN precursors. Unfortunately,rather sophisticated and expensive pitch pretreaments must be applied to a petroleum or coal tar pitch in order to produce a high performance fiber. The pitch softening temperature is much higher than PAN; thus stabilization can potentially be carried out quickly at higher temperatures. Nevertheless, stabilization of both PAN and pitch materials is exothermic and, in order to avoid overheating, thermal runaway and decomposition of the precursor, a less than advantageous temperature of oxidation must be used. Future developments in these areas should therefore involve methods to increase the stabilization rate and the development of new precursor materials. Specific topics might include stabilization in thermally stable environments (fluidized beds, liquids, etc.) and the development of alternative synthetic precursor materials (polymers and/or pitches). Cost of the final component, while very sensitive to the cost of raw materials, also involves all of the design, fabrication and testing costs. All of these component costs must be tightly controlled if economically viable performance increases are to be realized. When using composites, it has been found to be vitally important to pay strict attention to detail design if a maximum expected weight saving is to be realized. Designers must optimize the total vehicle weight and not simply substitute a carbon composite for a metal one. It is rumored for instance that, due to overly conservative design (i.e., the use of metallic joining and fabrication techniques etc.), the resulting weight of some composite structures has, in the past, turned out to be as heavy as similar components built from aluminum. 9.14 CONCLUSION
Many people appear to believe that despite nearly thirty years of development, carbon fibers are still an evolving space age material. Until very recently, there was a production
198 Carbon fibers over-capacity in the carbon fiber industry. For example, in 1995 it was estimated that 10 000 000 kg of carbon fibers were sold from an estimated capacity of 16 000 000 kg. Exactly how accurate these estimates were is difficult to assess; however, it can be concluded that the market is small and cannot accommodate many producers. For this reason, many producers have seen fit to evaluate their position in the industry. For example, Table 9.3 is a listing of carbon fiber processors taken from Chapman and Hall's directory, 'Carbon and High Performance Fibers' which was published in 1991. In 1996, only the first eight of these were still producing significant commercial quantities of PAN-based carbon fibers. Table 9.4 is a similar listing for pitch-based fibers. In this case only the first five appear to be active.
At the present time, downsizing of the industry, increasing use of low cost fibers and the resurgence of orders for new commercial aircraft that now use increased quantities of carbon fiber has brought industrial capacity and market requirements closer together. Indeed, some fiber types are now difficult to obtain. Nevertheless, any major growth of the carbon fiber industry depends on the discovery of a method to produce fibers for one-half or one-third of the present projected large volume price and the development of new inexpensive fabrication methods for structures. These developments will initiate major new transportation based markets for the material. At the present price, however, the use of carbon fibers will always be limited to competitive performance driven applications.
Table 9.3 PAN-based tow manufacturers
Company
Country
Trade name
Akzo Carbon Fibers Inc. (Fortafil Fibers Inc.) Amoco Performance Products Inc. RK Carbon Fibres Limited Mitsubishi Rayon Co. Ltd. Soficar SA (Toray Industries Inc.) Toho Rayon Co. Ltd Toray Industries Inc. Zoltek Corporation Akzo NV (Fibres and Polymers Division)(Enka AG) Anglo-Soviet Materials Ltd Asani Kasel Carbon Fiber Co. Ltd BASF Structural Materials Inc.' Formosa Plastics* Hercules Advanced Materials and Systems Company2 Korea Steel Chemical Co. Ltd Nikkiso Co. Ltd* N.W. Chemical Power Co.* Sigri GmbH Textron Specialty Materials
USA USA UK Japan France Japan Japan USA Holland USSR Japan USA Taiwan USA Korea Japan China Germany USA
Fortafil Thomel RK Pyrofil Torayca F Besfight Torayca Panex Tenax Sapem Hi-Carbolon Celion
* Not available. Now Hexcel; * Now Toho.
?
Magnamite Kosca ?
?
Sigrafil Avcarb
References
199
Table 9.4 Pitch-based tow manufacturers
Company
Country
Trade name
Amoco Performance Products Inc. Mitsubishi Kasei Corporation Petoca Ltd (Kashima Oil Co. Ltd) Tonen Corporation Nippon Petrochemicals Co. Ltd Kawazaki Steel Co. Ltd Kobe Steel Ltd" Mitsubishi Oil Co. Ltd* Nippon Carbon Co. Ltd* Nippon Steel Co. Ltd" Osaka Gas Co. Ltd (Donac Ltd) Showa Shell Sekiyu
USA Japan Japan Japan Japan Japan Japan Japan Japan Japan Japan Japan
Thomel Dialead Carbonic Forca Granoc KMFC ? ? ? ?
Donacarbo-F Carbonexel
* Not available.
REFERENCES Aggarival, R.K. 1977, Carbon 15291. Ashton, J.E., Halpin J.C. and Petit, P.H. 1969.Primer on Composite Materials: Analysis, Westport, Conn.: Technomic Publishing Co. Bacon, R. 1973. Chemistry and Physics of Carbon 9:2. New York Marcel Dekker, Inc. Bathia, G., Fitzer, E. and Kompalik, D. 1984. International Carbon Conference, Bordeaux, France, Paris: Group Francais d'etude des carbon. Boncher, E.A., Cooper, R.N. and Everett, D.H. 1970, Carbon 8:597. Bright, A.A. and Singer, L.S. 1979. Carbon 17: 59. British Patent 1341 161,1973. British Patent 1 371 621, 1974. White. British Patent 2 071 702,1981. Toho Beslon. Brooks, J.D. and Taylor, G.H. 1965. Nature 206:697. Brooks, J.D. and Taylor, G.H. 1968. Chem. Phys. Carbon 4243. Carbon and High Performance Fibers Directory, 1991. Chapman and Hall. Chieu, T.C., Timp, G., Dresselhaus, M.S., Endo, M. and Moore, A.W. 1983. Phys. Rev. B 27:3686. Chupalov, V.S., Migunov, E.I., Panov, V.P. and Tereshchenko, L.E. 1983. Zh. Prikl. Kkim. (USSR) 56:2595-2597. Corten, H.T. 1967. Modern Composite Materials (ed. Broughtman L.J. and Krock, R.H.). Reading, MA: Addison-Wesley. Clarke, A.J. and Bailey, J.E. 1973. Nature 243: 146-150. Daniel, I.M. and Ishai, 0. 1994. Engineering Mechanics of Composite Materials, Oxford University Press.
Deurberque, A., Ph.D. 1990. Thesis, Universite de Pau et des pays de l'adour. Deurberque, A. and Oberlin, A. 1991. Carbon 29: 621. Donnet, J.B. and Bansal, R. 1984. Carbon Fibers, Vol. 3. New York: Marcel Dekker, Inc. Edie, D.D. 1990. Carbon Fibers, Filaments and Composites. (eds. Figneiredo, J. Bernard, C.A. Baker. R.T.K. and Hiittenger, K.J.) Kluwer Academic, pp. 647-655. Edie, D.D. and Diefendorf, R.J. 1993. Carbon Fiber Manufacturing. Park Ridge, NJ: Noyes Publications. Erlemenko, I.N., Plyublines I. and Gulko, N.V. 1990, Chemical Modified Carbon Fibers. New York: VCH Publishers, Inc. Eshback, O.W. and Souders, M., Handbook of Engineering Fundamentals, New York: John Wiley. Ezekiel, H.M. 1969. Appl. Polym. Symp. 9: 315. Fitzer, E., Muller, K. and Schaeffer, W. 1971. Chemistry and Physics of Carbon, pp. 237-383. New York: Marcel Dekker, Inc. Fitzer, E. and Heine, M. 1988.Fibre Reinforcementsfor Composite Materials Vol. 2; pp. 73-148. Amsterdam: Elsevier. (ed. A. R. Bunsell). French Patent 2 328 723. National Research Development Corp. French Patent 1 535 800, 1968. North American Aviation, Inc. Glushchenko, V.M. and Griffen, L.A. 1982. Gibkie Elektroprovodnye Materialy I Ustroistva na Osnove dlya Obogreva Lyudei I Tekhniki, pp. 23-29. Grassia, N. and McGuchan, R. 1971a. Eur. Polymer J. 71091-1104.
200 Carbon fibers Grassia, N. and McGuchan, R. 1971b. Eur. Polymer 1. 7: 1357-1371. Griffith, A.A. P. 1920. Trans. R. Soc. A221:163. Guigon, M. 1985. Relations entre la microtexture et les proprietes mecaniques des fibres decarbone ex-PAN. D.Sc. Thesis (These d’Etat), Universite de Technologie de Compiegne, France. Guigon, M. and Oberlin, A. 1986a. Composites Sci. Technol. 25: 231. Guigon, M. and Oberlin, A. 198613. Composites Sci. Technol. 27: 1. Guigon, M., Oberlin, A. and Desarmot, G. 1984a. Fibre Sci. Technol. 20: 177. Guigon, M., Oberlin,A. and Desarmot, G. 1984b. Fibre Sci. Technol. 20: 55. Hadcock, R.N. 1982. Design and Analysis of Composite Structures. In Handbook of Composites. New York: Van Nostrand Reinhold co. Hamada, T., Nishida, S., Matsumoto , Y. and Endo, M. 1987.1,Mater. Res. 2: 850. Herring, H. W. 1966. NASA Rep. No. IND-3202. High Performance Composites. 1994. July/August. Jain, M.K. andAlhiraman,A.S. 1987.1.Mater. Sci. 22: 278. Jakubowski, J.J. and Subramanian, R.V. 1979. Chem. Abstr. 1981, 95: 98874. Johnson D.J. 1987. Chemistry and Physics of Carbon 20: 1.New York: Marcel Dekker, Inc. Johnson, J.W. 1969. A p p . Polymer Symp. 9: 229. Johnson, J.W., Potter, W., Rose, P.G. and Scott, G. 1972. Brit. Polymer I., 4: 527-540. Johnson, J.W., Rose, P.G. and Scott, G. 1970. Proc. 3rd Conf. lndustrial Carbon and Graphite, London: Academic Press, p. 443. Jones, R.M. 1975. Mechanics of Composite Materials, Washington, D.C: Scripta Book Co. Joseph, D. and Oberlin, A. 1983a. Carbon 21: 559. Joseph, D. and Oberlin, A. 1983b. Carbon 21: 565. Karpinos, D.M. and Izmalkov, O.M. 1982. Gibkie Elektroprovodnye Materialy I Ustroistva na Osnove dlya Obogreva Lyudei I Tekhniki. Kawamura, K. and Jenkins, G.M. 1970. A New Glassy Carbon Fiber. 1.Mater. Sci. 5: 262. Kowbel, W., Wapner, P. G. and Wright, M. A. 1988. 1.Phys. Chem. Solid 49: 11. Lafdi, K., Bonnamy, S. and Oberlin, A. 1991a. Mechanism of anisotropy occurrence in a pitch precursor of carbon fibres, Part I - Pitch A and 0. Carbon 29: 831. Lafdi, K., Bonnamy, S. and Oberlin, A. 1991b. Mechanism of anisotropy occurrence in a pitch precursor of carbon fibres, Part I1 - Pitch C.
Carbon 29: 849. Lafdi, K., Bonnamy, S. and Oberlin, A. 1991~. Mechanism of anisotropy occurrence in a pitch precursor of carbon fibres, Part I11 - Hot stage microscopy of pitch B and C. Carbon 29: 857. Lafdi, K., Bonnamy, S . and Oberlin, A. 1992. Textures and structures in heterogeneous pitch-based carbon fibres (as-spun, oxidized, carbonized and graphitized); comparison with homogeneous fibres. Carbon 31: 29. Lafdi, K. and Oberlin, A. 1994a.A tentative to characterize and elaborate anisotropic pitches and derived carbon fibres. Part I: preparation by separation. Carbon 32: 11. Lafdi, K. and Oberlin, A. 1994b.A tentative to characterize and elaborate anisotropic pitches and derived carbon fibres. Part 11: preparation by bubbling. Carbon 32: 61. Lavin, J.G. 1992. Carbon 30: 351. Levit, R.M. 1986. Conducting Synthetic Fibers, Khimiya, Moscow. Lewis, I.C. 1982. Carbon 20: 519. Lewis, I.C. and Nazem, F.F. 1987a. 18th Conference Carbon, Extended Abstracts, American Carbon Society, p. 190. Lewis, I.C. and Nazem, F.F. 1987b. 18th Conference Carbon, Extended Abstracts, American Carbon Society, p. 290. Mair, W.N. and Mansfield, E.H. 1987. William Watt 1912-1985. Biographical Memoirs of Fellows of the Royal Society, 33: 643-667. Matsumoto, T. 1985. Pure Appl. Sci. 57: 1553. Mochida, I., Shimizu, K., Korai, Y., Ohtsuka, H. and Fujiyama, 5.1988. Carbon 26: 843. Nazem, F.F. 1982. Carbon 20: 345. Nazem, F.F. and Lewis, I.C. 1986. Mol. Cryst., Liq. Cryst. 139: 195. Oberlin, A. and Guigon, M. 1988. Fibre Reinforcement for Composite Materials, (ed. A.R. Bunsell). Amsterdam: Elsevier, p. 149. Oberlin, A. and Oberlin, M. 1981. Revue Ckim. Miner. 18: 442. Oberlin A. 1984. Carbon 22: 521. Oberlin, A. and Guigon, M. 1984. Science and New Applicafions of Carbon Fibers, Toyohashi University of Technology, Japan. Ohtsuka, H. 1988. Mitsubishi Gas-Chemical Co. Kurashiki, Okayama 712, Japan. Reynolds, W.N. and Moreton, R. 1980. Philos. Trans. Roy. SOC.A294: 451. Reynolds, W.N. and Sharpe, J.V. 1974. Carbon 12: 103. Richter, E., Knoblauch, K., Juntgen, H. Deutsche Offen. Patent 3 412 761, 1984.
References Riggs, D.M., Shuford, R. and Lewis, R. 1982. Handbook of Composites. New York: Van Nostrand Reinhold. Riggs, D.M., 1979. Doctoral Thesis. Rensselaer Polytechnic Inst., Troy, New York. Robson, D., Assabghy , F.Y.I. and Ingram, D.J.E. 1972. J. Phys. D, 5: 169. Rosen, B.W. 1964. AIAA 2: 1985. Schwartz, M.M. 1984. Composite Materials Handbook. New York: McGraw-Hill. Shindo, A., Nakanishi, Y. and Soma, I. 1969. Appl. Polym. Symp. 9: 305. Singer, L.S. 1978. Carbon 16:409. Tanabe, Y., Yasuda, E., Machino, H. and Kimura. S. 1987. Ann. Mtg Jpn Ceramic Society, Nagoya, 77. Temovoi, K. S., Zemskov, V. S., Kolesnikov, E. B. and Mashkov, 0. A. 1985. Sorbitsionnaya v Khirurgicheskoi Klinike Defoksikatsiya (Detoxification Sorption in Surgery) Kishinev (USSR):Shtiintsa. Tsai, S.W. and Hahn, H.T. 1980. Introduction to Composite Materials, Technomic Publishing Co., Inc., Westport, CT. US Patent 3 107 152,1963. Ford and Mitchell. US Patent 3 533 741,1970. Courtaulds Limited. US Patent 3 627 466,1971. Steingiser. US Patent (12) 3 629 379, 1971. Otani. US Patent 3 650 668,1972. Celanese. US Patent 3 656 882,1972. Celanese. US Patent 3 656 883,1972. Celanese. US Patent 3 656 903,1972. Celanese. US Patent 3 660 140, 1972. Scola. US Patent 3 671 411,1972. Ray. US Patent 3 677 705,1976. Celanese. US Patent 3 723 150,1973b. Druin. US Patent 3 746 450,1973. Goan. US Patent 3 746 506,1973. Aitken. Druin. US Patent 3 754 957,1973~~ US Patent 3 767 773,1973. Turner. US Patent 3 767 774, 1973. Hou. US Patent 3 780 255,1973. Boom. US Patent 3 814 377,1974. Monsanto. US Patent 3 832 297,1974. Paul, Jr. US Patent 3 841 079,1974. Celanese US Patent 3 846 833,1975. Celanese. US Patent 3 859 187,1975. Druin.
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US Patent 3 894 884,1975. Druin. US Patent 3 917 776,1975. Mitsubishi Rayon. US Patent 3 919 383,1975. Singer. US Patent 3 972 984,1976. Nippon Carbon Co. US Patent (8) 3 974 264,1976. McHenry. US Patent (8) 3976729, 1976. Lewis, McHenry, Singer. US Patent 3 989 802,1976. Loo. US Patent 4 004 053,1976. US Patent 4 005 183,1977. Singer US Patent 4 009 248,1977. US Patent (4) 4 017 327, 1977. Lewis, McHenry, Singer. US Patent 4 024 227,1978. US Patent 4 031 288, 1978. Minnesota Mining and Manufacture Co. US Patent 4 039 341, 1976. National Research Development Corp. US Patent 4 069 297,1978. Toho Beslon Co. Ltd. US Patent (6) 4 208 267,1980. Diefendorf and Riggs. US Patent (5) 4 331 620,1982. Diefendorf and Eggs. US Patent 4 374 114,1983. Kim. US Patent (3) 4 376 747,1982. Nazem. US Patent 4 504 454,1985. Riggs. Watt, W. 1972. Carbon 10: 121. Watt, W. and Johnson, W. 1975. Mechanism of oxidation of polyacrylonitrile fibres. Nature 2 5 7 210-212. White, J.L. 1992. ONR Report for Contract No. 88-K-0424 and 89-J-3056. Wright, M.A., 1989. NASA Conference Publication 3054: 17. Wright, M.A. and Iannuzzi, EA. 1973. J. Comp. Mat., 7: 430. Wright, M.A. and Wills, J.L. 1974. J. Mech. Phys. Sol. 22: 161. Wright, M.A. and Palmer, K.R. 1994. Research into Structural Carbons, Materials Technology Center Publication, SIUC, Carbondale, Illinois, 62901. Yanagida, K., Noda, M., Sasaki, T. and Tate, K. 1991. 20th Conf Carbon, Extended Abstracts, American Carbon Society, p. 160. Yoneshoga, I. and Teranishi, H. 1970. Japanese Patent Specification 2774/70. Yooh, S.H., Korai, Y. and Mochida, I. 1994. Carbon 32: 281.
ORGANIC FIBERS
10
Linda L. Clements
A different type of h g h performance organic fiber, extended chain polyethylene fibers, was Before the first aramid fibers were introduced added in the 1970s. While inferior to inorganic in the 1960s and 1970s, organic fibers were relfibers in some properties, organic fibers provide atively low performance materials, primarily combinations of properties not available with used in textile applications. Now several different types of high performance organic inorganic fibers and so have made possible new fibers exist, all competitive with inorganic designs and applications. In this chapter, only high performance fibers in some or even most of their properties. organic fibers which are commercially availThe market demand for these fibers exceeds able will be discussed in detail, although fibers one billion dollars (Adams and Farrow, 1993a). which are nearing commercialization will be The main applications for high perfordiscussed briefly. For a more complete review mance organic fibers today are in asbestos of both commercially available and experireplacement, ballistics, rubber reinforcement, mental high performance organic fibers, see ropes and cables and composites. Most of the Yang (1989, 1992). usage is of aramid fibers, with over 18000 metric tons used each year. Both usage and existing capacity for other organic fibers are 10.2 ARAMID FIBERS only a fraction of this value (Adams and Farrow, 1993a). 10.2.1 OVERVIEW Tlus broad market for organic fibers is a direct outgrowth of applying the basic princi- Aramid fiber is the generic term for a specific ples of polymer science to produce a new and type of ’aromatic polyamide fiber.’ The US exceptional engineering material. In the 1950sit Federal Trade Commission defines an aramid was recognized that if a means could be found fiber as ‘a manufactured fiber in which the to form certain intractable polymers into fiber-forming substance is a long-chain synextended chain fibers, very high stiffnesses, thetic polyamide in which at least 85% of the strengths and use temperatures could be amide linkages are attached directly to two achieved. The difficulty of producing such aromatic rings.’ Thus, in an aramid, most of the amide fibers was solved in the 1960sby spinning from groups are directly connected to two aromatic liquid crystalline solutions. The first fibers prorings, with nothing else intervening. It should duced by t h s process were the aramids, which have since been followed by other such fibers. not be surprising that aramids have quite different properties from nylons and other conventional polyamides since the latter polymers contain few if any aromatic groups in the Handbook of Composites. Edited by S.T. Peters. Published main chain of the polymer. in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7 10.1 INTRODUCTION
Aramid fibers
203
Aramid fibers can be separated into two include DuPont’s KevlarO, Akzo’s TwarorP, types: the para- aramids and the meta-aramids. Teijin’s TechnoraO and Kaiser VIAM‘s Amosa In para-aramids, the chain-extending bonds are and S W @fibers, while meta-aramids include in the para-position on the aromatic ring, as in DuPont’s Nomexs and Teijin’s TeijinconexO poly-p-phenylene terephthalamide (PPTA) (Fig. fibers. Hoechst AG also markets a para-aramid 10.1(a)),co-poly-p-phenylene/3,4’-oxydipheny- fiber in Europe. The para-aramids are the fibers lene terephthalamide (Fig. lO.l(b)) and used in high performance applicationsand thus poly-p-phenylene-benzimidazole-terephthala- will be emphasized in this chapter. mide (Fig. lO.l(c)). In meta-aramids, on the other hand, the chain-extending bonds are in 10.2.2 MANUFACTURE the meta-position on the aromatic ring, as in poly-m-phenyleneisophthalamide (MPIA) (Fig. Historically, meta-aramid fibers were the first 10.1(d)). Commercially available para-aramids to be produced, with DuPont’s Nomex fiber 0
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\ H NI
d+0
Fig. 10.1 Structural formulae of (a) the para-aramid poly-p-phenylene terephthalamide (PPTA), (b) the para-aramid co-poly-p-phenylene/3,4’-oxydiphenyleneterephthalamide, (c) the para-aramid poly-pphenylene-benzimidazole-terephthalamide(PBIA), and (d) the meta-aramid poly-m-phenylene isophthalamide (MPIA).
204 Organicfibers being introduced in the 1960s. The first paraaramids were synthesized in 1965 by S.L. Kwolek of DuPont (Kwolek, 1971; Kwolek, 1972; Kwolek, 1974).Forming these into usable fibers is very difficult because para-aramids show no melting point and are soluble in a limited number of solvents. The problem of spinning the polymer into fibers was solved for PPTA following the discovery that the polymer would dissolve in strong acids to form a liquid crystalline solution. Undiluted sulfuric acid is the solvent usually used. Blades (1973, 1974) devised a special manufacturing process - known as continuous dry jet wet spinning - for forming the liquid crystalline solution into filaments. The polymer solution is extruded through spinnerets at elevated temperature through an air layer into a coagulating water bath. The cold water bath also contains a base to neutralize and remove the retained acid. Continuous dry jet wet spinning is the manufacturing technique used for most para-aramid fibers. Teijin’s Technora fiber, however, is produced by wet spinning followed by drawing (Hongu and Phillips, 1990).
10.2.3 STRUCTURE
The excellent properties of para-aramids result from both chemistry and physical microstructures. In both meta- and para-aramids, the aromatic rings in the backbone chain produce high thermal resistance. In addition, in paraaramids the orientation of the chain-extending bonds produces a polymer which is an extended-chain rigid rod. Spinning produces a fiber made up of extended-chain crystallites which are almost completely aligned parallel to the draw direction and to each other. The crystallites have a very high length-to-diameter ratio and extensive interconnection of molecules between crystallites. Thus, an unbroken ’infinite’ filament can be formed. Within the crystallite the chains are bonded to one another by hydrogen bonds, as shown in Fig. 10.2. Although these bonds are not nearly as strong as the covalent bonds which occur within the molecules, hundreds or even thousands of such bonds form between adjacent para-aramid molecules. Since the molecules are rigid, the only way to separate them in tension is to break all of the hydrogen bonds at once. This requires a large force and
0 I
H
H
H
A&=(-*
0
I
I
0
-N
II
\ llc - e
-C
\
H
H
\
0
0
-0 H
Fig. 10.2 Schematic showing hydrogen bonding between PPTA molecules in the crystallite.
Aramid fibers 205 is the reason para-aramid fibers are exceptionally strong in axial tension. However, since the bonds can be broken easily one at a time, the fibers are quite susceptible to damage by bending, buckling or transverse loading. In meta-aramids, on the other hand, a crooked chain results. Since even in pure tension the chain-extending bonds can flex and rotate, meta-aramids are much less rigid than para-aramids and not as strong. However, because the chains are more flexible, metaaramids are easier to manufacture than para-aramids and are less expensive.
mechanical, thermal, physical and other properties. This anisotropy may produce design limitations, but can also be used to advantage.
Physical and thermal properties
Table 10.1 compares the physical and thermal properties of some representative aramid fibers. Due to their highly aromatic and ordered structure, aramids have very high thermal resistance for organic materials. They do not melt prior to decomposition,in spite of the fact that they are technically classified as thermoplastics. This is because melting of the 10.2.4 PROPERTIES crystalline phase, like rupturing the fiber in Aramid fibers offer some significant advan- tension, would require that all of the hydrogen tages over other fibers, but also have their bonds between two molecules be severed at drawbacks and limitations. Both advantages once. Nonetheless, because of decomposition, and limitations will be described more fully in their temperature resistance is not equal to the sections on properties and in the sections that of inorganic fibers. Thermogravimetric on design considerations and applications. analysis of Kevlar fibers shows that weight Both DuPont’s Kevlar family of fibers and loss begins at above 350°C (660°F)in air (Penn Akzo‘s Twaron fibers are based upon PPTA and Larsen, 1979; Yang, 1992), with complete (Fig. lO.l(a)). Teijin’s Technora fiber and the decomposition occurring at between 427 and para-aramid marketed by Hoechst AG in 482°C (800 and 900°F) (DuPont, 1992a). Europe, on the other hand, are a para-aramid Exposure to elevated temperature will copolymer, co-poly-p-phenylene/3,4’-oxy- degrade the properties of aramid fibers. Figure diphenylene terephthalamide (Fig. lO.l(b)).It is 10.3 shows the strength retention of Kevlar 29 likely that Kaiser VIAM’s SVM fibers are poly- and Technora fibers as a function of time and p-phenylene-benzimidazole-terephthalamide temperature. This change in properties occurs (PBIA), (Fig. lO.l(c)) rather than PPTA as a result of slow oxidation. For this reason, (Gerzeski, 1989). Kaiser VIAM’s Armos fiber the long-term use temperature of para-aramid may be PBIA or PPTA. Both DuPont’s Nomex fibers is typically limited to about 150-175°C and Teijin’s Teijinconex fibers are based upon (300-350°F). MPIA (Fig. lO.l(d)).These chemical and strucIn the transverse direction para-aramids are tural differences produce different properties like most other materials in that they expand for the fibers. In addition, differences in spin- with increasing temperature. However, in the ning conditions and, most importantly, longitudinal direction the fibers actually conpost-spinning heat treatments are used to alter tract somewhat as temperature increases. The properties further. For example, by changing negative thermal expansion coefficient of processing conditions, Kevlar fibers can be para-aramids can be used to advantage to produced with elastic moduli ranging from 63 design composites with tailored or zero therto 143 GPa (9 to 21 Msi) and elongations at mal expansion coefficient. Aramids are flame resistant but can be break from 1.5 to 4.4%. Because of the anisotropy of their microstruc- ignited. While pulp or dust of Kevlar may conture, para-aramid fibers have very anisotropic tinue to smolder once ignited, fabrics do not
206 Organicfibers Table 10.1 Physical and thermal properties of representative aramid fibers
Fiber
Kevlar 49
Twaron HM
Technora
Nomex
Teijinconex
Type
para-aramid
para-aramid
para-aramid copolymer
meta-aramid
meta-aramid
DuPont 1992a 1.44 (0.0520)
Akzo 1990,1991 1.45 (0.0524)
Teijin 1989,1993 1.39 (0.0502)
DuPont 1981,1993g 1.38 (0.0499)
Teijin 1991 1.38 (0.0499)
-538°C" (1000°F)
>5OO0C (>932"F)
-
>371"C (>700"F)
Decomposition temperature in air
427482°C (800-900°F)
500°C (930°F)
500°C (930°F)
371°C (700°F)
400430°C (750-805°F)
Long-term use temperature in air
149-177°C (300-350°F)
Longitudinal linear thermal expansion coefficientb 10-6/ "C
-4.9 (-2.7)
Reference for data Density g cm-3 (lb in-?) Melting temperature
-3.5 (-1.95)
-6.2 (-3.4)
+15 (+8.3)
+20 (+11)
1.09 (0.26)
1.21 (0.29)
1.05 (0.25)
0.13 (22)
0.13 (22)
OF)
Transverse linear thermal expansion coefficientb: /"C ( / O F )
+66" (+37)
Specific heatb kJ/kg K (BTU/lb OF)
1.42 (0.34)
1.42 (0.339)
Longitudinal thermal conductivityb W/m K BTU/h ft OF)
4.11' (2.38)
4.0 (2.3)
Transverse thermal conductivityb W/m K BTU/h ft OF)
4.82' (2.79)
5.0 (2.9)
3.5%
3.5%
2.0%
4.5%
5.0-5.5%
Typical filament diameter pm in)
12 or 15 (0.48 or 0.59)
12 (0.48)
12 (0.48)
max: 15-17 (0.6-0.7)
-10 to 15 X 45 (-0.4 to 0.6 X 1.1)
Typical filament shape
round
round
round
oval to dogbone
oval to dogbone
Equilibrium moisture contentb
* Data from Yang, 1992.
Varies with temperature; room temperature values are given. Data from Chiao and Chiao, 1982.
Arumidfibers
rL
I
I
I
I
207
100
.
be
80
C
0
+ C
2w
60
L11
fm
e
I=
40
-
iz 20 -
0' 0.1
- Technorag I
I
I
I
1
10
100
1000
L
Time, h
Fig. 10.3 Strength retention of Kevlar 29 and Technora fibers following elevated temperature exposure (DuPont, 1992a; Teijin, 1989).
continue to burn when the flame source is removed (DuPont, 1992a). The lower thermal conductivity of aramids compared to inorganic fibers can improve the fire resistance of their composites, since aramids do not readily conduct heat into the more volatile matrix.
Mechanical properties Composite materials are most commonly used because of their superior strength and/or stiffness at a given weight as compared to conventional structural materials. Figure 10.4 compares the specific strengths and specific stiffnesses of various reinforcing fibers. (The strengths and stiffness in Fig. 10.4 are expressed in units of grams per denier (gpd). This is a textile term often used for organic fibers which measures specific strength and/or stiffness. This term is further explained in the appendix to this chapter.) As can be seen, aramid fibers perfonn very well. In fact, until the emergence of high strength intermediate modulus carbon fibers and the commercialization of polyethylene
fibers in the mid-l980s, aramid fiber composites had the highest specific strengths of all composite materials. Although composites from newer fibers have taken over that position, aramids still offer outstanding combinations of properties, such as high specific strength, toughness, creep resistance and moderate cost, for specific applications. Table 10.2 compares the mechanical properties in axial tension of several commercially available aramid fibers. Aramid fibers have some definite limitations. They are weak in bending and show obvious damage if subjected to kinking or buckling. As a result, they are also weak in compression (where microbuckling is inevitable) and in transverse tension (wherebond-by-bond breakage of hydrogen bonds is likely). In addition, even though the para-aramid chain is quite polar in nature, almost all of the polar groups are fully involved in hydrogen bonding to other aramid molecules. As a result, paraaramid fibers do not form strong bonds with other materials such as composite matrices,
208 Organicfibers Table 10.2 Axial tensile mechanical properties of representative aramid fibers Fiber
Reference
Spec$c gravity
Initial tensile modulus, GPa (Msi) Bare"
Epoxyimpregnatedb
Tensile strength, MPa ( h i ) Bare"
Bare" 3.6
Kevlar Type 956, 1500 denier
DuPont, 1993h
1.44
71.8 (10.4)
Kevlar 29 Type 964, 1500 denier
DuPont, 1992a, DuPont, 19938
1.44
70.5 (10.2)
83.0 (12.0)
2920 (424)
Kevlar 49 Type 965, 1140 denier
DuPont, 1992a, DuPont, 1993g
1.44
112.4 (16.3)
124.0 (18.0)
3000 (435)
Kevlar 68 Type 9898, 1420 denier
DuPont, 1993g
1.44
99.13 (14.4)
3050 (442)
-
2.9
Kevlar 119 1500 denier
DuPont, 1990
-
54.6 (7.9)
3050 (442)
-
4.4
Kevlar 129 denier unspecified
DuPont, 1993i
1.44
96.0 (13.9)
3380 (490)
-
3.3
Kevlar 149 Type 965A, 1140 denier
DuPont, 19938
1.47
142.7 (20.7)
2340 (339)
-
1.5
Kevlar HT Type 964C, 1000 denier
DuPont, 19938
1.44
99.1 (14.4)
3370 (489)
-
3.3
Kevlar KM2 850 denier
DuPont, 1992d
-
63.4 (9.2)
3280 (476)
-
4.0
Twaron
Akzo, 1991
1.44
70 (10.2)
2800 (406)
Twaron Perkins, 1993 Type 2000, 930 denier 'microfilament'
1.44
88 (12.8)
3230 (468)
Twaron HM
Akzo, 1991
1.45
Armos 58.8 tex
Kaiser VIAM,
103 (14.9) 147 (21.4)
SVM 58.8(300) X17-1000
Kaiser VIAM, 1993g' Gerzeski, 1989
-
1993a'
1.43
2920 (424)
Epoxyimpregnatedb
Elongation at break, %
-
3600 (525)
3500 (508) -
3500 (508) -
3.6
3.6 3.3
2.5 3.2
123 (17.8) Continued on next page
Aramid fibers 209 Table 10.2 Continued
Fiber
Reference
Specific gravity
Initial tensile modulus, GPa (Msij Bare"
Epoxyimpregnatedb ~~
-
Teijin, 1989
1.39
Nomex Type 430, 1200 denier
DuPont, 1993g
1.38
11.6 (1.68)
-
Teijinconex
Teijin, 1991
1.38
7.9-9.7 (1.1-1.4)
-
11.6-12.2 (1.7-1.8)
-
a
Teijin, 1991
1.38
Bare"
Elongation at break, %
Epoxyimpregnatedb
-
3440 (498)
-
4.6
596 (86.6)
-
28.0
610-670 (88-97)
-
3545
730-850 (110-120)
-
20-30
Data for DuPont fibers taken from conditioned yarns tested according to ASTM Standard D885. Modulus data for Akzo fibers from testing according to ASTM Standard D885M. Test technique unspecified for Akzo fiber strengths and elongations and for all data from Kaiser VIAM and Teijin fibers. Data for DuPont fibers taken from epoxy-impregnated strands tested according to ASTM Standard D2343. Data for Akzo fibers from testing according to impregnated strand test method DIN 65356, part 2. Test technique unspecified for Kaiser VIAM fibers. Preliminary data. 50
40
. . . . I
..
I . I . . . I l l l l l l . . . .
-
l l . . I I I I . . I . I . .
-
PED HM Armos 0
Carbon TlOOOG 30 - Technora 0
'pectra 'Oo0
0
Dyneema SKBO
Carbon
Tekmilon 'spectra 900
. Vectran HS 0 Kevlar 49 . Twaron O O S ~0 Carbon T-300 . .waron HM 2o - S-Glass 0 Kevlar 149 0 Carbon T-50
.
PE (H.C.)
Boron
E-Glass Carbon P-100 .
Steel
Bare"
~
73 (10.6)
Technora
Teijinconex HT
Tensile strength, MPa (ksij
1
2 10 Organicfibers and/or prepared using other fabrication processes, the general trend is valid: aramid fiber composites have poor off-axis properties. In axial tension, both aramid fibers and their composites are linear to failure. In spite of this fact, the same microstructural characteristics which lead to the weakness of aramid fibers in buckling also make them very tough. During failure, the widespread bending, buckling and other internal damage to the fibers absorbs a great deal of energy. Similarly, the strength of aramid fibers is not very strain rate sensitive: an increase in strain rate of more than four orders of magnitude only decreases the tensile strength by about 15%. (Abbott et al., 1975) This property alone provides design advantages over all inorganic and many other organic fibers. The mechanical properties of aramid fibers decrease with increasing temperature. Figure 10.5 shows the fiber elastic modulus as a function of temperature for several organic fibers. At 177°C (350°F) the modulus of para-aramid fibers is about 80% of that at room temperature. Figure 10.6 compares the fiber tensile
further aggravating the poor transverse, bending and compressive properties of the fiber itself. The basic chemical structure differences between the aramid fibers produce many of the mechanical property differences seen in Table 10.2. The ether (-0-) linkages in the backbone of the Technora copolymer fiber produce a lower modulus than that of Kevlar and Twaron PPTA-based fibers. On the other hand, the additional cyclic ring in the SVM PBIAbased fibers produces a higher basic modulus. However, heat treatment and other fabrication steps can also alter mechanical properties significantly, as is seen in the property differences between the various Kevlar fibers. The mechanical properties of aramid composites are illustrated in the data of Table 10.3. For this filament-wound composite the longitudinal compressive strength was about one-eighth that in longitudinal tension, the inplane shear strength was one-seventy-fifth and the transverse tensile strength over two hundred times smaller. While the relative values of properties may change for composites made from other aramid fibers and/or other matrices
Table 10.3 Mechanical properties of a filament-wound composite of 60 vol YO aramid fiber in a room-temperature curable epoxy matrix (Clements and Moore, 1977)
Fiber: DuPonf’s Kevlar 49, Type 968, 1420 denier Matrix: 100 parts Dow Chemical DER 332 (diglycidyl ether of bisphenol-A epoxy) and 45 parts Jefferson Chemical reffamine T-403 polyether triamine 1 day at room tnnperuture, postcure 16 h ut 85°C (185°F) Cure: Elastic constants: Longitudinal Young’s modulus E,,, GPa (Msi) Transverse Young’s modulus E,,, GPa (Msi) Shear modulus G,,, GPa (Msi) Major Poisson’s ratio vl, Minor Poisson‘s ration u,,
Ultimates:
81.8 f 1.5” (11.9 k 0.22) 5.10 k 0.10 (0.74 & 0.014) 1.82 k 0.09 (0.26 f 0.013) 0.310 k 0.035 0.0193 f 0.0014
Tension
Longitudinal strength, MPa (ksi) 1850 f 50 (268 f 7.3) 2.23 f 0.06 Longitudinal ultimate strain, Yo Transverse strength, MPa (ksi) 7.9 k 1.1(1.15 f 0.15) 0.161 f 0.023 Transverse ultimate strain, Yo Shear stress at 0.2%, offset, MPa (ksi) Shear strain at 0.2% offset, Yo
Compression
In-plane shear
235 f 3 (34.1 k 0.4) 0.48 5 0.3 53 f 3 (7.7 f 0.4) 1.41 f 0.12
-
-
24.2 f 2.4 (3.51 k 0.35) 1.55 k 0.16
-
a
Limits are 95% confidence limits. Each value is the result of five or more tests.
Aramidfibers
211
s
si - 5 25
-
1992a; Teijin, 1989).
3500
-1
a
-
1
8
-
Techno ra
m
q
: -
2500 1
Kevlar
8 =
;:0003 n g 2000 0,
t?
1500
3i
500
400
300
-
-
.-
Y"
f m c
200 ?!
Polyester
looo;, Nylon
- 100
500 0
, ,: 0
Fig. 10.6 Tensile strength as a function of temperature for two para-aramid fibers and for two polymer fibers and steel (DuPont, 1993h; Teijin, 1989).
212
Organicfibers
strength as a function of temperature for sev- wet transverse tensile and in-plane shear eral organic fibers. For Kevlar fiber the strengths were only about half of the 52% r.h. strength at 177°C (350°F) is about 80% of that values. The data in boiling water illustrate that at room temperature, while for Technora the the drops in strength due to the presence of strength is about 70% of the room temperature moisture alone were almost as severe as those value. On the other hand, at cryogenic tem- due to the combined presence of moisture and peratures modulus increases slightly and elevated temperature. This relative loss in properties is less for the Technora para-aramid strength is not degraded. The presence of moisture also reduces the co-polymer fiber. Care must be exercised when mechanical properties of aramid fibers and using aramid composites in high moisture their composites. The effect upon longitudinal applications. Both para-aramid co-polymers and homotensile properties is relatively small, but the loss is pronounced for off-axis properties. polymers exhibit very little creep. In general, Table 10.4 illustrates this loss for Kevlar 49 creep strain increases with increasing temperafiber in a room-temperature curable epoxy. ture, increasing stress and decreasing fiber The longitudinal tensile strength in water at modulus. Like all high performance fibers, room temperature was 88% of that for com- under long term loading, para- aramids are posites equilibrated at room temperature and subject to stress rupture, i.e. failure of the fiber 52% relative humidity (r.h.). The wet longitu- under sustained loading with little or no dinal compressive strength, on the other hand, accompanying creep. Figure 10.7 compares the was only 75% of the 52% r.h. value, while the stress rupture performance of Kevlar 49 to that Table 10.4 The effect of environments on the mechanical properties of a filament-wound composite of 50 vol Yo of an aramid fiber in a room-temperature curable epoxy matrix (Wu, 1980) Fiber:
DuPont's Kevlar 49, 4560 denier
Matrix: 100 parts Dow Chemical DER 332 (diglycidyl ether of bisphenol-A epoxy) and 45 parts Jefierson Chemical Jeffamine T-403 polyether triamine Cure:
Infrared heating, postcure 2 h at 100°C (212°F) Strength, MPa (ksi)
________
23"C, dry
23°C' 52% r.k.
23"C, water
1OO"C, water
Longitudinal tension
1370 f 6 2 " (199 f 9)
1340 f 112 (194 f 16)
1190 f 62 (173 f 9)
1150 f 1 2 4 (167 f 18)
Longitudinal compression
188 f 1 2 (27.3 f 1.7)
169 f 20 (24.5 f 2.9)
126 f 22 (18.3 f 3.2)
107 f 2 1 (15.5 f 3.0)
Transverse tension
7.6 f 1.6 (1.10 f 0.23)
74 f 1.2 (1.07 k 0.17)
3.9 f 0.7 (0.57 f 0.10)
3.6 f 0.2 (0.52 f 0.03)
Transverse compression
31.3 f 3 . 2 (4.54 kO.46)
29 f 4.0 (4.21 f 0.58)
22.5 f 3 . 2 (3.26 f 0.46)
22.1 f 23.6 (3.20 f 3.42)
In-plane shear
27 f 3 . 0 (3.92 f 0.44)
26.5 f 1.6 (3.84 f 0.23)
13.8 f 2.2 (2.00 f 0.32)
13.6 f 2.5 (1.97 f 0.36)
-
4.1
7.8
8.9
Hygrothermal Properties Equilibrium moisture concentration, Yo ~
Limits are 95% confidence limits. Each strength is the average of five tests.
Aramid fibers 213
0
+, 100
! 2 .I4
Kevlafl
90
49
1 W
o
80
d
.
a a
70
u
,"
60
a
a 0
50
.rl
rl
2a
40
10-2
10-1
1
io
io2
103
104
105
Lifetime, h
Fig. 10.7 Stress-rupture behavior of epoxy-impregnated Kevlar 49 fibers compared to that of epoxyimpregnated S-glass fibers (Chiao,Chiao and Sherry, 1976).
of Sglass. Para-aramids perform well under these conditions, but the phenomenon of stress rupture must be considered in any design where long term loading is anticipated. Strength retention cannot be used to estimate the remaining life of aramid fibers or composites under long term load (Chiao, Sherry and Chiao, 1976),so estimates of long term behavior must be derived from actual data, or accelerated testing methods (Chiao and Chiao, 1982). Para-aramid fibers and their composites perform very well in fatigue. For aramids, tension-tension fatigue generally is not of significant concern in applications where an adequate static safety factor has been used (Yang, 1992). Aramid composites have been found to be superior to glass fiber composites in both tensile-tensile and flexural fatigue loading. For the same lifetime (cycles to failure), Kevlar 49/epoxy composites can operate at a significantly larger percentage of their static strength than can glass-reinforced composites
(DuPont, 1986). Para-aramids also can be expected to perform better than carbon fibers in fatigue (Teijin, 1989; Yang, 1992). Technora para-aramid co-polymer is found to have even better fatigue resistance than the para-aramid homopolymer fibers (Teijin, 1989).
Chemical and environmental properties PPTA fibers are quite stable chemically; their resistance to neutral chemicals is usually very high. They are, however, subject to attack by acids and bases, especially by strong acids. Because the spin process used for Teijin's Technora para-aramid co-polymer produces a very pure polymer, the chemical and environmental resistance of Technora is superior to that of the PPTA fibers. Table 10.5 reports the resistance of Kevlar and Technora fibers to various chemicals. Technora has better acid and alkali resistance than PPTA and its steam resistance is also superior.
214 Organicfibers Table 10.5 Stability of para-aramid fibers in various chemicals
Concentration, Temperature,
None Slight Moderate Appreciable Degraded
40 40 90 90 20 10 10 10 10 10 10 20 40
21 (70) 95-99 (203-210) 21 (70) 95-99 (203-210) 20 (68) 71 (160) 20-21 (68-70) 20-21 (68-70) 21 (70) 99 (210) 99 (210) 95 (203) 95 (203)
1000 100 Tb 100 K 100 100 T 10 100 T 100 K,T 1000 100 10 100 T 100 T
28 10 10 saturated saturated
21 (70) 21 (70) 95-99 (203-210) 95 (203) 180 (356)
1000 1000 100 100 15
Formic Hydrochloric Nitric Phosphoric Sulfuric
Alkalis Ammonium hydroxide Sodium hydroxide
Organic solvents Acetone Benzene
100 100 100 Carbon tetrachloride 100 Ethylene chloride 100 Ethylene glycol/water 50/50 Ethylene glycol 100 Gasoline 100 Gasoline-leaded 100 Methyl alcohol 100 N-Methyl pyrrolidone 100
3 10 10 Sea water 100 Sea water (New Jersey) 100 Steam 100 100 100 100 Water, tap 100
a
hr
"C ( O F )
Acids Acetic
Other Sodium chloride
Effect on breaking strength"
Yo
Chemical
Portland cement
Time,
K
T
K K K K K
K T
K
T T
100 K 784 T 1000 K 100 1000 T 1000 T 300 784 T 1000 K 1000 K,T 100
21 (70) 99 (210) 121 (250) 95 (203)
1000 100 100 1000 1 yr 400 48
K K
100
T
100 100
K
-
K
K
boil 20 (68) 21 (70) boil 20 (68) 99 (210) 95 (203) 20 (68) 21 (70) 21 (70) 95 (203)
120 (248) 150 (302) 150 (302) 200 (392) 99 (210)
Kb
K K
T
K T K T K T
None, 0-10% strength loss; slight, 11-20% strength loss; moderate, 2140% strength loss; appreciable, 41430% strength loss; degraded, 81-100% strength loss. K is for Kevlar aramid fiber (DuPont 1989,1993h);T is for Technora aramid fiber (Teijin, 1989).
Aramid fibers Para-aramids are strong ultraviolet (W) absorbers. Upon exposure, the yellow or gold fibers turn first orange and then brown, due to degradation. The degradation occurs only in the presence of oxygen and is not enhanced by either moisture or atmospheric contaminants (DuPont, 1992a). Extended exposure may cause a loss of mechanical properties. Bare 1667 dtex (1500 denier) Kevlar 29 was found to have 71% strength retention after 1 month of outdoor exposure in Wilmington, DE and 43% after 4 months (Yang, 1992).In both processing and applications, para-aramids must be protected from W exposure, such as by painting or coating. However, since para-aramids are self-screening, UV protection may also be effected simply by dense packing of the fiber itself, with or without a matrix. Thus, bare 12.7mm (0.5 in) 3-strand Kevlar 49 rope was found to have 90% strength retention after 6 months outdoors in Florida and 69% strength retention after 24 months (DuPont, 1986). Unlike inorganic fibers, aramid fibers absorb water. For some aramid fibers the equilibrium moisture content (see appendix on page 241 for defirution) is quite high (5% for SVM, 7% for Kevlar, Kevlar 29 and Twaron), moderate for others (3.5% for Kevlar 49 and Twaron HM) and reasonably low for some (2% for Armos and Technora and about 1% for Kevlar 149) (Akzo, 1991; Kaiser VIAM, 1993a; Teijin, 1989; Yang, 1992). The equilibrium moisture content is directly proportional to the relative humidity, rising for Kevlar 49 to 6.2% at 96% r.h. (DuPont, 1992a). Absorbed moisture has only a small effect upon the tensile properties of the fibers, but a significant effect upon the transverse tensile, compressive, shear and flexural properties of the composite. The gain of moisture is completely reversible and once removed produces no permanent property changes. Electrical and optical properties Aramid fibers are electrical insulators. The process used to make the Technora fiber, however, leaves it with fewer ionic impurities than
215
in the para-aramid homopolymers and thus improved electrical properties. Technora fiber has a resistivity of 5 x lo’* Q/cm (Teijin, 1989). The dielectric constant of PPTA is 3.85 (Allied, 1989). The refractive index of Kevlar 49 fiber is 2.0 parallel to the fiber axis and 1.6 perpendicular (DuPont, 1986).Aramid fibers are opaque and are yellow to gold in color. 10.2.5 TREATMENTS
Unlike inorganic fibers, few surface treatments are used on aramid fibers to promote matrix adhesion. One reason is the futility of increasing the matrix bonding to the surface of a fiber which readily fails by defibrillation. Most dramatic improvements in fiber/matrix bonding give only modest improvements in off-axis strengths since they simply move the locus of failure from the surface to the interior of the filament. In other cases, longitudinal tensile strengths are adversely affected by otherwise successful surface treatments. Not all attempts at designing surface treatments have been unsuccessful, but for the most part the surface treatment used on commercial fibers is minimal compared to that used for inorganic fibers. Finishes - lubricants which aid in subsequent processing steps - are applied to aramid fibers for some applications. Available finishes are designed for such purposes as lubrication during weaving operations, improving abrasion resistance for cable applications or better performance in rubber goods. If the fiber is to be used in a high performance composite, however, the user will usually wish to avoid or remove any finish before impregnating the fiber with a matrix. Commercial aramid fibers may also be twisted. Twist may be quite useful in some applications and a small amount of twist will increase the strength of bare yarn or cord. [This optimum twist for Kevlar fibers occurs at a twist multiplier of 1.1.At about this value, the strength of bare yarn is the highest and the modulus is only slightly decreased from the
2 16 Organic fibers
untwisted level (DuPont, 1992b).] Twist will make the fiber easier to handle, make subsequent weaving or braiding operations easier and will improve the abrasion resistance of the fiber. It is also required for rope and cable applications. However, once the fiber is used in composite matrix fiber, twist is not desirable. This is because twist interferes with full impregnation of the fiber with resin and with stress transfer between adjacent fiber bundles. It also increases stress concentrations, particularly at higher twist levels. For this reason, most of the aramid fiber manufacturers supply most or all of their fibers untwisted or with minimal twist. 10.2.6 FORMS AND AVAILABILITY
Table 10.6 lists most of the commercial types of para-aramid fibers. Some of these fibers are readily available in a variety of fiber deniers, package sizes, finishes and so forth, while others are available only in limited quantities for specific applications. Due to constant changes in market conditions and other factors, the user is advised to check with the fiber manufacturer concerning current availabilities. The mechanical properties of fibers with different deniers and/or finishes and other treatments will vary somewhat from each other and from the nominal values given in Table 10.2. In addition to the yarns, tows and rovings listed in Table 10.6, DuPont’s Kevlar fibers are also available as staple (short fibers), floc (precision cut fibers of very short lengths) pulp (very short and high fibrillated fibers) and in specialty compounded forms (DuPont, 1992a, 1993h).A variety of fabrics are also produced. In addition, DuPont produces a colored fiber, Kevlar 100, in sage green, yellow, black and royal blue (Yang, 1992).Both Teijin’s Technora and Akzo’s Twaron fibers are available as staple and chopped fiber (very short lengths) and in a variety of fabrics (Teijin, 1989;Akzo, 1991). Technora is also marketed in black as well as natural color. Kaiser VIAM‘s Armos and SVM fibers are also expected to be offered as tape,
staple, pulp and in various fabrics. While the meta-aramid fibers are not usually used as fiber reinforcements in composites, they are used extensively as reinforcements for honeycomb sandwich core materials. The use of such materials along with composite face sheet panels has greatly extended the overall usage of composite materials, particularly in the aerospace industry. Information about the availability and package sizes of the fibers shown in Table 10.6, about other products and about special formulations can be obtained from Table 10.7. At the time of this publication, Kaiser VIAM’s Armos and SVM fibers are just being imported from Russia. For this reason, information on the fibers and their availability is limited in this chapter, but should be readily available later from the contact given in Table 10.7. Pricing Para-aramid fibers are currently priced from about $20 per pound for the larger denier fibers to about $60 per pound for most of the small denier, higher modulus fibers. (However, some of the very fine denier specialty fibers from some manufacturers cost hundreds of dollars per pound.) Prices can vary significantly for similar fibers of different deniers or from different manufacturers and thus price quotes should always be obtained before any decision is made upon use of a specific fiber. 10.2.7 DESIGN CONSIDERATIONS
In the 1970s and early 1980s aramids began to replace carbon and glass fibers in many applications. However, the development of high strength intermediate modulus carbon fibers in the mid-1980s and the commercialization of tough, high strength polyethylene fibers reversed this trend. Today aramid fibers are used mainly in applications where they offer a unique combination of properties, such as high specific strength combined with toughness and creep resistance.
Aramidfibers 217 Table 10.6 Availability of commercial para-aramid fibers"
Product (reference)
Count dtex (den)
Filament
~ _ _ _ _ _ _ _ _
number/ diameter yarn pm ( 1 C 3 i n )
Kevlarb (DuPont, 1993g; Yang, 1992) 666 Type 950 1110 (1000) 1670 (1500) 1000 1000 2500 (2250) 3330 (3000) 1333 Type956
800 (720) 1110 (1000) 1670 (1500) 2500 (2250) 3330 (3000)
Comments/typical applications
490 666 1000 1000 1333
12 12 15 15
(0.48) (0.48) (0.59) (0.59)
Finish: tire reinforcement
12 12 12 15 15
(0.48) (0.48) (0.48) (0.59) (0.59)
Mechanical rubber goods: hoses, belts, etc.
Kevlar 29 (DuPont, 1992b,1993a, 1993b, 1993f, 19938) Type 960 1670 (1500) 1000 12 (0.48) Cordage finish: high lubricity for improved 3330 (3000) 1333 15 (0.59) abrasion resistance; ropes and cables 100OOR' 12 (0.48) 17 OOO(15 000) 1110 (1000) 1670 (1500) 3330 (3000) 5000 (4500) 17 OOO(15 000)
666 1000 1333 3000 lOOOOR
Type 962
1670 (1500) 3330 (3000)
1000 1333
12 (0.48) 15 (0.59)
No finish; ropes and cables
Type 963
3330 (3000)
1333
15 (0.59)
Textile finish; non-apparel ballistic armor
Type964
215 (200) 430 (400) 1110 (1000) 1670 (1500)
134 267 666 1000
12 12 12 12
(0.48) (0.48) (0.48) (0.48)
Textile finish; ballistics and apparel, ignition cables
Kevlar 49 (DuPont, 1992b, 1993c-g) Type 965 61 (55) 25 215 (195) 134 420 (380) 267 1270 (1140) 768 1580 (1420) 1000 2400 (2160) 1000
15 12 12 12 12 15
(0.59) (0.48) (0.48) (0.48) (0.48) (0.59)
Textile finish; woven reinforcement in aerospace composites, ballistic armor, and printed circuit boards
12 12 12 12 15 15 15 12 12
(0.48) (0.48) (0.48) (0.48) (0.59) (0.59) (0.59) (0.48) (0.48)
No finish; marine composites, fiber optic cable reinforcement, ropes, filament-wound composites
Type968
215 (195) 420 (380) 1270 (1140) 1580 (1420) 2400 (2160) 3160 (2840) 4800 (4320) 5070 (4560) 7900 (7100)
134 267 768 1000 1000 1333 2000R 3200R 5000R
12 12 15 12 12
(0.48) (0.48) (0.59) (0.48) (0.48)
Textile finish; ropes and cables
Type 961
Continued on next page
218 Organicfibers Table 10.6 Continued
Product (reference)
Count dtex (den)
Filament
Commenfs/fypicalapplications
number/ diameter yarn pm ( 1 P i n ) ~
_
_
Type 978
1580 (1420) 2400 (2160) 5070 (4560)
1000 1000 3200R
12 (0.48) 15 (0.59) 12 (0.48)
Cordage finish: high lubricity for improved abrasion resistance; ropes and cables
Type 989
1580 (1420) 2400 (2160) 3160 (2840) 4800 (4320) 6300 (5680) 7900 (7100) 9500 (8520)
1000 1000 1333 2000R 2666R 5000R 4000R
12 15 15 15 15 12 15
Textile finish; fiber optic cable reinforcement
Kevlar 68 (DuPont, 1992b, 1992c, 19938) 215 (195) 90 Type9568 1000 1580 (1420) Type9898
420 (380) 1580 (1420) 2400 (2160) 3160 (2840) 4800 (4320) 7900 (7100)
267 1000 1000 1333 2000 5000R
(0.48) (0.59) (0.59) (0.59) (0.59) (0.48) (0.59)
12 (0.48) 12 (0.48)
High performance mechanical rubber goods
12 12 15 15 15 12
Textile finish; fiber optic cable reinforcement
(0.48) (0.48) (0.59) (0.59) (0.59) (0.48)
Kevlar 129 (DuPont, 1990, 1993h) Type 956E 1670 (1500) 1000
12 (0.48)
Power transmission belts, high-performance tires, high fatigue applications
Kevlar 129 (DuPont, 1992c, 1993h, 1993i) 666 Type 956C 1110 (1000)
12 (0.48)
Mechanical rubber goods
12 12 12 12
Personal body armor
Type964C
830 (750) 930 (840) 1110 (1000) 1580 (1420)
500 6OOL 666 1000
Kevlar 249 (DuPont, 1992c, 19938) Type965A 420 (380) 267 768 1270 (1140) 1000 1580 (1420) Type 968A
1270 (1140) 1580 (1420) 4730 (4260) 7890 (7100)
768 1000 3000R 5000R
(0.48) (0.48) (0.48) (0.48)
12 (0.48) 12 (0.48) 12 (0.48)
Woven reinforcement in aerospace composites, hard ballistic armor, printed circuit boards
12 12 12 12
No finish; marine composites, fiber optic cable reinforcements, ropes, filament-wound composites
(0.48) (0.48) (0.48) (0.48)
Kevlar HT (DuPont, 19938) Type 964C 1110 (1000)
666
12 (0.48)
Advanced ballistic protection
Kevlar K M 2 (DuPont, 1992d) 945 (850)
560
12 (0.48)
Ballistic protection: helmets, composite armor Continued on next page
Ararnidfibers 219 Table 10.6 Continued
Product (reference)
Count dtex (den)
Filameizf
Comments/typical applications
number/ diameter Fnz ( l C 3 i n ) yarn
Twarond (Akzo, 1990,1991; DeCos, 1993) Type 1000
420 (380) 840 (760) 1100 (990) 1260 (1130) 1680 (1510) 2520 (2270) 3360 (3020)
250 12 (0.48) 500 12 (0.48) 750 10.5 (0.41) 750 12 (0.48) 1000 12 (0.48) 1500R 12 (0.48) 2000R 12 (0.48)
Standard finish; multipurpose
Type1001
420' (380) 840 (760) 1100' (990) 1260 (1130) 1680 (1510) 3360 (3020)
250 12 (0.48) 500 12 (0.48) 750 10.5 (0.41) 750 12 (0.48) 1000 12 (0.48) 2000R 12 (0.48)
Adhesive-activated finish; tires, mechanical rubber goods, composites
Type 1010
1680 (1510) 3360 (3020)
1000 2000R
12 (0.48) 12 (0.48)
Very low finish level; composites
Type 1020
1680 (1510)
1000
12 (0.48)
Special finish for increased abrasion resistance; cables, ropes, nets
Type 1030 17 OOO(15 300)
5000R
12 (0.48)
PTFE + silicone oil impregnated; braided packings
Type 1031 14 OOO'(12 600)
5000R
Type1040
420 (380) 840 (760) 1100' (990) 1260 (1130) 1680 (1510)
250 500 750 750 1000
Type 1041
1260' (1130) 1680' (1510)
750 1000
12 (0.48) 12 (0.48)
Adhesive-activated finish; fabrics
Type2000
930 (840) 1110 (1000) 1680 (1510)
1000 1000 1000
6.6 (0.26) 8 (0.31) 12 (0.48)
Standard finish; high tenacity for ballistic applications. 930 dtex fiber is 'microfilament'.
Twaron HM (Akzo, 1991) Type 1055 1210 (1090) 1610 (1450) 2420 (2180) 3220 (2900) 4830 (4350) 6440 (5800) 8050 (7245) Type 1056
1210 (1090) 1610 (1450) 2420 (2180) 6440 (5800) 8050 (7245)
12 (0.48) 12 12 10.5 12 12
(0.48) (0.48) (0.41) (0.48) (0.48)
PTFE + silicone oil impregnated; braided packings Tangled yarn; multipurpose
750 1000 1500R 2000R 3000R 4000R 5000R
11.5 11.5 11.5 11.5 11.5 11.5 11.5
(0.45) (0.45) (0.45) (0.45) (0.45) (0.45) (0.45)
Standard finish; multipurpose
750 1000 1500R 4000R 5000R
11.5 11.5 11.5 11.5 11.5
(0.45) (0.45) (0.45) (0.45) (0.45)
Very low finish level; composites
Continued on next page
220 Organicfibers Table 10.6 Continued
Product (reference)
Count
Filament
Comments/typical applications
-
dtex (den)
number/ diameter yarn pm (1Win)
Twaron IM (Akzo, 1991) Type 1111
420' (380) 1260 (1130) 1680 (1510) 2520 (2270)
250 750 1000 1500R
12 12 12 12
(0.48) (0.48) (0.48) (0.48)
Easily removed finish; fiber optic cable reinforcements, ballistics, composites
-
-
-
Twisted 48 t/m; multipurpose
-
-
-
-
-
-
-
Type A1 lubricating finish Lubricating finish on 'acidic' fiber Lubricating finish on 'neutral' fiber Lubricating finish on 'acidic' fiber Type A1 lubricating finish on 'acidic' fiber Two different heat treatments available Type A1 lubricating finish on 'acidic' fiber
Armos (Kaiser VIAM, 1993a) 588 (530) SVM(Kaiser VIAM, 1993b-j) 63 (57) 143 (130) 294 (265) 294 (265) 294 (265) 588 (530) 588 (530)
-
-
-
-
Technoru' (Teijin, 1989; Mahn 1993) T-200
1110 (1000) 1670 (1500)
666 1000
12 (0.48) 12 (0.48)
Rubber reinforcement
T-202
440 (400) 1670 (1500)
1667 1000
12 (0.48) 12 (0.48)
Rubber reinforcement, pre-activated type
T-220
1110 (1000) 1670 (1500)
666 1000
12 (0.48) 12 (0.48)
Rope, cable, and cord
T-221
1110 (1000) 1670 (1500)
666 1000
12 (0.48) 12 (0.48)
Rope, cable, and cord
T-230
1670 (1500)
1000
12 (0.48)
Fiber-reinforced plastics, rope
T-240
60 (55) 110 (100) 220 (200) 440 (400) 1110 (1000)
36 67 133 267 666
12 12 12 12 12
(0.48) (0.48) (0.48) (0.48) (0.48)
Woven and knitted fabrics, fiber-reinforced plastics
T-241
1670 (1500) 8330 (7500)
1000 5000R
12 (0.48) 12 (0.48)
Woven and knitted fabrics, fiber-reinforced plastics
T-360
608 (55) 220 (200) 440 (400) 1110 (1000) 1670 (1500)
36 67 133 267 1000
12 12 12 12 12
'Spunnized' yarn (made up of long but not continuous filaments) for protective clothing and other fabric applications
T-370
220 (200) 440 (400)
133 267
For footnotes see next page
(0.48) (0.48) (0.48) (0.48) (0.48)
12 (0.48) 12 (0.48)
High tenacity 'spunnized' yarn for reinforcement of rubber, etc.
Aramid fibers 221 Table 10.7 Sources of information o n commercial aramid fibers
Information source
Product Armos and SVM fibers
Kaiser VIAM; 880 Doolittle Drive, San Leandro, CA 94577, USA
Kevlar fibers
DuPont Fibers; P.O. Box 80705, Wilmington, DE 19880-0705, USA, (800)
4-KEVLAR Technora fibers
Teijin Limited, 11, 1-chome, Minamihonmachi, Chuo-ku, Osaka 541, Japan Teijin America Inc; 10 East 50th Street, New York, NY 10022, USA
Twaron fibers
Akzo, Aramide Maatschappij v.o.f., P.O. Box 9300,6800 SB A r h e m , Westervoortsedijk 73, The Netherlands Akzo Fibers Inc., 801-F Blacklawn Rd., Conyers, GA 30207, USA
The outstanding toughness of aramids is often the reason they are used over cheaper, stiffer or even stronger fibers. Unlike glass and carbon composites, aramid composites loaded in compression, flexure or shear fail in a non-brittle manner, with significant work being required to fail the composite. Their fatigue resistance is also excellent. If other concerns such as cost or stiffness preclude the use of aramid composites, aramids are often used as a hybrid with another fiber to improve the toughness of the composite. The poor off-axis and compressive properties of aramid fibers must be considered in any design. However, because of their high strength in axial tension and their toughness, aramid fibers are often outstanding in applications such as pressure vessels where the loading is almost totally in longitudinal tension. Aramid fibers absorb moisture. Where either the physical swelling of the fiber or the amount of moisture absorbed is of significant concern, one of the lower absorption aramids, such as Kevlar 149, Armos, or Technora should be considered.
Aramids are strong UV absorbers and deteriorate when exposed to ultraviolet light. Protective coatings or the self-screening ability of the fiber should be used to avoid deterioration. Aramid fibers are opaque and thus the penetration of resin into the fiber bundles cannot be determined visually for a aramid composite as it can for those made with glass fibers. In fabric applications the weave used is important to the resulting properties. The same is true for sandwich construction. In these cases, the fiber, fabric, or honeycomb supplier can provide design assistance. The choice of resin system for use with aramid fibers is an important one. Epoxy resins give better translation of fiber properties than do polyesters, producing better shear strength and flexural properties, but lower impact resistance. Vinyl ester resins give both good shear strength and impact resistance. Thermoplastic matrices are also used, particularly in chopped fiber composites, because of their improved impact resistance over thermosets. However,
Footnotes for Table 10.6 All availabilities are subject to market conditions and should be verified with the manufacturer. All Kevlar fibers are supplied untwisted. E The ' R indicates that this fiber is a 'roving,' meaning in this case that it is composed of more than one 'end' of yam. Twaron fibers are normally supplied untwisted. In some circumstances twist may be supplied on special request. e Under development. Technora fibers are supplied with twist as requested. Finishes are supplied as requested or as is appropriate to the application. For special applications,Technora fibers can be supplied in larger than 12 pm filament diameters. These fibers, and others 'spun' yarns (composed of discontinuous filaments) are normally measured by '(English) cotton count' (ECC) rather than dtex or denier, where ECC = 5315/denier. a
'
222 Organicfibers Many of these are not as structural composites. For example, aramids are used in many rope and cable applications. In mooring ropes to secure oil tankers and to anchor off-shore oil platforms, the lighter weight compared to steel makes the aramid ropes much easier to handle. In addition, they do not corrode, are easier to maintain and have an extension under load which is far superior to both steel and other organic fibers. Aramids are widely used to reinforce mechanical rubber goods. The largest volume of such usage is in pneumatic tires, where aramids are lighter than steel and offer higher strength and modulus than other organic fibers. Significant usage is also seen in belts and hoses. The excellent fatigue and creep resistance of aramids are important factors in their usage in these applications. Corrosion resistance and electrical resistivity may also be important. Aramids are also used in athletic shoes and in rubberized sheet materials as used in aircraft evacuation slides and life rafts. In some cases, non-composite applications have led to composite uses. For example, aramids have long been used in soft body armor, where the fibers absorb and disperse bullet impact energy to other fibers in the fabric weave. This application has now seen a derivative usage in rigid composite ballistic armor, composite helmets and composite spa11 liners. In these applications the toughness, RF transparency and fire and corrosion resistance of aramid fibers were significant factors in their selection. In spite of significantly higher fiber costs than glass, aramids are used in canoes, kayaks, racing shells and small boats where maximizing strength and minimizing weight are important. Aramids offer weight savings for superior speed and better handling and/or improved range and fuel economy. Toughness and overall durability and vibrational damping are also superior with aramids. The 10.2.8 APPLICATIONS superior properties of aramids allow boats to Aramid fibers are used in numerous applica- be built at an overall cost only 10-15% higher tions, some of which are listed in Table 10.6. than with glass fibers and with superior perfor-
for thermoplastics the penetration of the resin into the fiber bundle and the quality of the fiber-matrix bond is almost always of concern. Because aramids are very tough fibers, they are somewhat difficult to cut and their composites can be difficult to machine. Special shears and other tools are available for cutting aramids and many successful machining techniques have been developed. The fiber manufacturers are an excellent source of information in this area. As with all high performance fibers, aramids should be handled with care before and during processing. Rough handling will damage any high performance fiber. In addition, because of their sensitivity to ultraviolet light, aramids should be protected from such exposure. The fibers also should not be exposed to excessive moisture prior to processing. If the fiber is to be twisted, braided, or woven, it is preferable to condition the fiber for one to two days at room temperature and intermediate moisture content prior to processing (DuPont, 1993h). However, if the fiber is to be resin-impregnated and processed directly into a composite, so long as fiber handling is careful, superior properties may be attained by drying the fiber prior to processing. Tlus is because of improved bonding of resin to the filament surfaces. Aramid fibers present minimal safety or environmental concerns. In lifelong animal inhalation studies with Kevlar fibers, no health effects were observed at any workplace levels. Nonetheless, as with any textile fiber, inhalation of fibrous particles should be avoided. Extensive animal and human skin patch tests with Kevlar fibers have shown no sensitivity and little irritation, and rat feeding studies have shown oral toxicity to be very low. Combustion by-products are similar to wool. Aramid yarns are also essentially inert in the environment (DuPont, 1993h).
Extended
h i i i
polyefhylcviefibers 223
mance (DuPont, 1983). These same properties properties but they also have limitations that have led to the use of aramids in skis. must be considered in design. Their high strength-to-weight ratio comCommercially available high strength, h g h bined with outstanding toughness has led to modulus polyethylene fibers include Spectrakh' numerous applications of aramids in aero- fibers from Allied-Signal Corporation, space. In both civilian and military aircraft, the DyneemaO SK60 from Dyneema Vof, Tekmilon" toughness of aramids - and resulting resis- from Mitsui Petrochemicals and a new, as yet tance to damage from impacts ranging from unnamed, fiber from Hoechst Celanese. bird strikes to shrapnel - insures their continued usage. Engine nacelles and the tail cone on 10.3.2 MANUFACTURE the McDonnell Douglas DC-9-80 are made from Kevlar composites and approximately The traditional method of producing fibers 10% of the empty airframe weight of De from polyethylene is to spin them from a polyHavilland Aircraft's DASH-8 turboprop com- mer melt. This technique yields fibers muter aircraft is Kevlar composite. Aramid composed of folded-chain crystalline regions composites are also widely used in rotorcraft with non-crystalline regions interspersed. With and other vertical lift aircraft. extraordinary means, the modulus of the absolute best of such fibers can be brought to about 80 GPa (11.5Msi). It was long recognized, 10.2.9 CONCLUSIONS however, that if polyethylene could somehow Although composites of other fibers have now be produced with extended chain crystallinity, supplanted aramid composites as having the a very high modulus fiber would result. [The highest specific strengths, aramids still offer theoretical modulus for polyethylene is combinations of properties not available with 320 GPa (46 Msi) (Adams and Eby, 1987).] any other fiber. For example, aramids offer Following earlier work by Pennings, in the high specific strength, toughness and creep late 1970s Smith and Lemstra of DSM (The resistance, combined with moderate cost. Netherlands) developed a process with comHowever, the applications of aramid compos- mercial potential which yielded a highly ites continue to be limited by their poor oriented extended-chain polyethylene fiber compressive and off-axis properties and in (Hongu and Phillips, 1990).At the same time, some applications, their tendency to absorb both Toyobo Inc. of Japan and Allied Chemical water. Nonetheless, aramids will continue to Company in the USA were working on a simbe the fiber of choice where properties such as ilar approach. DSM, however, was the first to outstanding impact resistance combined with patent the process and both Toyobo and Allied judged it impossible to circumvent the basic creep resistance are critical. patent filed by DSM. Thus, both companies entered into technical association with DSM to 10.3 EXTENDED CHAIN POLYETHYLENE produce polyethylene fibers. Toyobo Inc. FIBERS linked with DSM to form the joint venture Dyneema Vof - to produce and market the 10.3.1 OVERVIEW new fiber. In the USA, Allied-Signal is licensed High performance polyethylene fibers, with from DSM/Stamicarbon to produce and maroutstanding strength-to-weight and stiffness- ket a similar fiber. The process which is used to produce most to-weight performance, show promise in various specialized applications. While such commercial high strength, high modulus polyfibers are not as widely known as aramid and ethylene fibers is called gel spinning, the name carbon fibers, they possess many superior derived from the gel-like appearance of the
224 Organicfibers as-spun and quenched fibers. Ultra-high molecular weight linear polyethylene is dissolved in a volatile solvent to form a dilute isotropic solution that is then spun through a spinneret and quenched in cold water to form a gel precursor fiber. Following solvent extraction, this fiber is then hot-drawn to a very high draw ratio (= 30), yielding a very highly oriented, highly crystalline, lightweight fiber (Dyneema, 1987; Jaffe, 1989; Ward and McIntyre, 1986; Yang, 1992). Another approach to producing a high strength polyethylene fiber is melt extrusion followed by multiple stage drawing of a much lower molecular weight polyethylene. The modulus of an experimental fiber of this type, 220 GPa (32 Msi), is the highest ever achieved for polyethylene (Adams and Eby, 1987). The new polyethylene fiber from Hoechst Celanese is the only commercial version of such a fiber. This fiber has only about 50% of the strength and 75% of the modulus of gel spun fibers. In this case, the expense of dealing with a volatile and potentially toxic solvent is avoided, lowering the overall price of the fiber significantly. 10.3.3 STRUCTURE
Fig. 10.8 Schematic illustrating the difference between (a) conventional polyethylene fibers and (b) gel-spun extended chain fibers.
Figure 10.8 illustrates the difference between conventional polyethylene fibers and gel-spun hydrogen bonds nor strong covalent bonds or melt-extruded and drawn extended-chain between them. They are, in fact, held together fibers. Figure 10.8(a)is a schematic of a con- by weak dispersion-type van der Waals bonds ventional melt-drawn polyethylene fiber. The which have a distinct effect upon properties. fiber consists of folded chain crystallites, mostly oriented in the draw direction, which 10.3.4 PROPERTIES are joined to one another by tie molecules and have between them interspersed non-crys- Polyethylene fibers offer a unique combinatalline material. Figure 10.8(b)is a schematic of tion of properties: low specific gravity, high a gel-spun and hot-drawn extended-chain specific modulus, high specific strength, high polyethylene fiber. Such fibers show minimal energy to break, high abrasion resistance, chain folding, high crystallinity and a very excellent chemical resistance, good ultraviolet resistance and low moisture absorption. They high degree of axial orientation (>95%). Since these fibers are based on polyethyl- have outstanding anti-ballistic and vibrational ene, they have a density of only two-thirds damping characteristics, as well as a low that of aramid fibers and about one-half that of dielectric constant. However there are tradecarbon fibers. However, the polyethylene crys- offs involved in the use of polyethylene fibers. tallites have neither relatively strong They are limited to fairly low use temperatures,
Extended chain polyethylene fibers 225 they produce composites with poor off-axis treated Spectra 900 fiber in an epoxy matrix and compressive properties and have poor was found to be -9 x lO"/OC (-5 x 104/OF) in creep resistance. the axial direction and 100 x lO"/"C (56 x As with aramid fibers, the anisotropy of 104/"F) in the transverse direction. The axial their microstructure gives polyethylene fibers thermal expansion coefficient of a similar comanisotropic mechanical, thermal and physical posite of Spectra 1000 fiber was -10 x lO"/"C properties which can be used to advantage in (-5.6 x 10"/OF) and the transverse coefficient some applications. was 105 x lO"/OC (58 x 10"/"F) (Allied, 1989). Polyethylene fibers are the only high performance fibers with a specific gravity of less than Physical and thermal properties 1 and thus are the only fibers that float. Their Polyethylene fibers have a relatively low melt- density is about two-thirds that of aramid ing point [147"C (297"F)I and thus a low use fibers and about half that of carbon fibers. temperature. In general, polyethylene fibers Polyethylene fibers will burn slowly if ignited, are limited to use below 100°C (212°F). They decomposing into carbon dioxide and water. will, however, tolerate brief exposure (30 min The filament diameters of commercial polyor less) at temperatures near the melting point ethylene fibers are relatively large, typically without major property loss (Dyneema, 1987; 23-38 pm (0.91-1.50 x in), although the Weedon and Tam, 1986). diameter of Mitsui's Tekmilon monofilament As would be expected from the lower melt- fibers can be as large as 121 pm (4.76 x in). ing temperature, the properties of polyethylene The filament cross-section is typically irregufibers are much more sensitive to temperature lar and somewhat elliptical. than are aramids. Like aramid fibers, polyethylene fibers contract with temperature Mechanical properties in the axial direction, while expanding in the transverse direction. The thermal expansion Gel-spun polyethylene fibers offer some coefficient of a composite of 60 vol% plasma- tremendous advantages over other fibers. As Table 10.8 Axial tensile mechanical properties of representative high performance polyethylene fibers
Fiber
Reference
Fiber type
Spec@ gravity
Tensile modulus, GPa (Msi)
Tensile strength, MPa (ksi)
Elongation at break, %
-
Dyneema SK60
Dyneema, 1987 gel-spun
0.97
87 (12.7)
2620 (380)
-
Hoechst Celanese fiber
Hoechst meltCelanese, 1993 extruded
0.96
1300 (189)
4
Spectra 900
Allied, 1993
gel-spun
0.97
55 (8.0) 86-103 (12.5-14.9)
2080-2400 (300-350)
3.6-3.7
Spectra 1000
Allied, 1993
gel-spun
0.97
128-171 (18.6-24.8)
2740-3000 (397435)
2.8-3.1
Tekmilon monofilament
Mitsui, 1989
gel-spun
0.96
59-98 (8.6-14.2)
1470-3430 (213498)
4-6
Tekmilon multifilament
Mitsui, 1989
gel-spun
0.96
88.3 (12.8)
2450 (356)
3
226 Organicfibers
can be seen in Fig. 10.4, these fibers offer very high specific stiffnesses and specific strengths, equivalent or superior to all of the aramid fibers and to most of the carbon fibers. This superior performance is offered at a lower price than that of competitive fibers. Table 10.8 compares the mechanical properties of representative commercially available polyethylene fibers. Like aramid fibers and for similar reasons, polyethylene fibers have poor compressive and off-axis properties. Since the fiber is held together internally by only very weak van der Waals bonds, the transverse strength of the fiber is even worse than that for the aramids. In addition, the inertness of the polyethylene fiber means that the untreated fiber bonds
very poorly to a matrix. Although gas plasma surface treatment can improve the interfacial bond strength significantly, polyethylene fiber composites will still have poor off-axis properties. Table 10.9 gives mechanical properties for Spectra fiber composites, including those made from plasma-treated fibers. In spite their weak transverse strength, but because of the non-stick nature of polyethylene and thus its low coefficient of friction, polyethylene fibers perform much better than aramids in abrasion resistance and polyethylene fabrics are much less easily damaged than are those of aramid fibers. The abrasion resistance of polyethylene fibers can be up to ten times that of aramids (Dyneema, 1987)and can be improved
Table 10.9 Mechanical properties of Spectra polyethylene fiber composites” (Allied, 1989)
Matrix: Bisphenol A based epoxy __
-
Spectra 900 Axial Tensile Properties: Volume percent fiber Modulus, GPa (Msi) Strength, MPa (ksi) Elongation, % Axial Compressive Properties: Volume percent fiber Modulus, GPa (Msi) Strength, MPa (ksi) Elongation, YO Flexural Properties: Volume percent fiber Modulus, GPa (Msi) Strength, MPa (ksi) Short Beam Shear Properties: Volume percent fiber Strength, MPa (ksi) a
58 27 f 1 (4.0 f 0.1) 552 zk 90 (80 f 13) -
70 32 f 5 (4.7 k 0.7) 52 k 2 (7.5 f 0.3) -
~
Spectra 900 P T ~ Spectra 1000
Spectra 100 PT
-
54 50 f 3 (7.2 f 0.5) 889 f 55 (129 f 8) 2.1 i-0.4
50 24 f 1 (3.5 f 0.1) 676 f 103 (98 15) 3.6 f 0.2
53 50 (7.3) 1034 c 228 (150 f 33)
70 40 k 5 (5.8 + 0.7) 59 + 1 (8.6 c 0.2)
55 19c6 (2.7 f 0.9) 72 f 3 (10.5 f 0.4) 3.8 k 0.5
65 54 f 3 (7.8 f 0.4) 69 f 1 (10.0 f 0.2) 3.8 f 0.2
*
-
-
58 22 k 1 (3.2 f 0.2) 145 f 7 (21 k 1)
54 30f1 (4.3 f 0.2) 200 f 7 (29 f 1)
54 23 zk 1 (3.3 k 0.2) 159 +. 7 (23 f 1)
53 38 f 3 (5.5 f 0.5) 214 f 7 (31 k 1)
58 8.3 f 0.7 (1.2 f 0.1)
54 28.3 f 0.7 (4.1 f 0.1)
54 9.0 c 0.7 (1.3 f 0.1)
53 21 e 3 (3.1 f 0.4)
Numbers of specimens tested and criteria for limits not specified. PT indicates a fiber with gas plasma surface treatment.
Extended chain polyethylene fibers 227 even further by the use of lubricants. Because of their high strength, polyethylene fibers exhibit very high energy to break. On a per-weight basis, the impact energy absorption of polyethylene composites is superior to that of all other fiber composites. Polyethylene fibers are more affected by temperature than are higher melting point fibers. The loss in modulus as function of temperature is shown in Fig. 10.9 for Tekmilon multifilament fiber and Spectra fibers. Fig. 10.10 shows the loss in strength as a function of temperature for Tekmilon multifilament, Spectra 900 and Spectra 1000 fibers. Because of their very high specific strength at room temperature, however, polyethylene fibers still outperform most other fibers to about 100°C (212°F). Room temperature strength retention of polyethylene fibers following annealing at temperatures of up to 125°C (260°F) is excellent, while modulus loss following such
r L 1 . 8 8
100
0 a
I
,
I
I
I
I
,
exposure is 20-30%. The loss in both modulus and strength are reduced if annealing is performed under tensile loading. Unlike aramid fibers and their composites, polyethylene fibers and composites show very little or no loss of properties, axial or off-axis, when exposed to moisture. Creep resistance of extended-chain polyethylene is of concern. Because of its low melting temperature, the resistance of the fiber to creep, even at room temperature, is less than ideal. This is significant, since the creep of carbon, glass and aramid fibers is minimal. Spectra 1000 is a 'stabilized' version of the fiber, which shows better creep resistance than the Spectra 900 fiber. Figure 10.11 shows the creep response of the two Spectra fibers at room, elevated and low temperatures. At low load levels at room temperature and/or at low temperatures the creep encountered is not severe, especially for the Spectra 1000 fiber, but at higher loads or temperatures the creep is much
,
I
S
*
- 15
-
.-ul
75 -
- 10
c)
i
ul
-33
= -3
50 -
U
0
z
I
- 5 25
0
-
~
~
.
l
l
~
l
.
~
'
~
~
~
0
~
l
~
~
~
~
I
Fig. 10.9 Modulus as a function of temperature for Spectra 900, Spectra 1000, and Tekmilon multifilament polyethylene fibers (Prevorsek, 1989; Mitsui, 1989).
*
l
l
l
l
228 Organicfibers Temperature,
200
150
100
50
OF
250
300 500
-
3000
- 400
-
2500 0
%
-
2000 1
.-
300 -$
f0,
f
VI
5
1500
1
1000
-
500
-
C
i7l
50
25
0
100
75
Temperature,
125
e
-
200 3;
-
100
0 150
OC
Fig. 10.10 Strength as a function of temperature for Spectra 900, Spectra 1000, and Tekmilon multifilament polyethylene fibers (Allied, 1991e; Mitsui, 1989).
-
Spectra 900
.___-0
------RT, 10% Load
~
0
~
10
20
30
40
50
60
70
80
90
100
time, h
Fig. 10.11 Creep of Spectra extended chain polyethylene fibers (a) at room temperature and 10% of static ultimate and at room temperature and 30% of static ultimate.
Extended chain polyethylene fibers 229
6:
.-C
e
t
- Spectra 900 .___ Spectra 1000
G 1
0
25
50
75
100
time,
125
150
175
h
(a)
time,
h
Fig. 10.11 (Continued) Creep of Spectra extended chain polyethylene fibers (b) at 5°C (41°F)and 20% of static ultimate, and (c) at 70°C (160°F)and 275 MPa (40 ksi), which is 18% of static ultimate for Spectra 900 and 11%of static ultimate for Spectra 1000 (Allied, 1991a, 1991b, 1991c, 1991d).
230
Organic fibers
more significant. The creep of polyethylene fibers does not preclude their use in applications such as sailcloth or structural reinforcement, but does require that the creep demands of an application be carefully evaluated. Because of their relatively poor creep resistance, polyethylene fibers are often hybridized with other, more creep resistant fibers in applications where prolonged loading is anticipated. The fatigue resistance of polyethylene fibers is excellent. In one test of loading and unloading of ropes, polyethylene fiber ropes withstood approximately eight times the cycles before break as aramid fiber ropes (Weedon and Tam, 2986). T ~ indicates E a superiority in tensile fatigue even to aramid fibers, which are known for their excellent fatigue resistance.
Chemical and environmental properties Polyethylene is inert. It is stable in almost all organic solvents and in a variety of other chemicals. It is also biologically inert. It is the best of all high modulus fibers in an alkaline environ-
ment (Jaffe, 1989). It shows superior chemical resistance to PPTA in hydrochloric, nitric and sulfuric acid (Dyneema, 1987).Table 10.10 compares the chemical resistance of Spectra polyethylene fiber to that of aramid fiber. Polyethylene fibers also show good resistance to UV exposure. After 100 hours UV exposure in a fadeometer, Dyneema SK60 retained 70% of its original strength and after 1500 hours, retained 25% strength. This latter exposure is equivalent to about 2 years of outdoor exposure (Dyneema, 1987). Polyethylene fibers are hydrophobic and thus absorb very little moisture. The moisture regain of polyethylene fibers is less than 1%. Their weatherability is excellent: after 600 hours exposure in a Weatherometer, Tekmilon fiber retained 80% of its strength and 90% of its modulus. Following similar exposure, an aramid fiber had only 40% strength retention (Mitsui, 1989).Because of their excellent chemical and moisture resistance, articles made with polyethylene fibers can be cleaned in soap and water.
Table 10.10 Comparison of strength retention after chemical immersion for polyethylene and aramid fibers (Allied, 1989)
Strength retention, YO Spectra Chemical Sea water Hydraulic fluid Kerosene GasoIine Toluene Glacial acetic acid 1M hydrochloric acid 5M sodium hydroxide Ammonium hydroxide (29%) Perchloroethylene 10"/0detergent solution Chlorine bleach Too weak to test
Aramid
6 months (4380 h)
2 years (17 500 h)
6 months (4380 h)
2 years (17 500 h)
100 100 100 100 100 100 100
100 100 100 100 100 100 100 100 100 100 100 73
100 100 100 93 72 82 40
98 87 97
100 100 100 100
91
42 70 75 91 0
a
Extended chain polyethylene fibers 231 Electrical and optical properties
bonding of the fiber to matrices. Through 1992, Allied-Signal marketed plasma-treated Polyethylene fibers are electrically non-conSpectra fibers. However, polyethylene surface ductive. The dielectric constant of Spectra treatments are available after fiber purchase fiber is 2.2, with a loss tangent of 2 x lo4 from specialty companies. The fiber manufac(Allied, 1993).This compares to dielectric conturers can suggest sources of these services. stants for aramid fibers of 3.85, quartz fibers of The inertness and high abrasion resistance 3.78 and E-glass fiber of 6.31 (Allied, 1989). of polyethylene fibers means that little or no They are white in color and are transparent to finish is required. This is fortuitous since adheX-rays, radar and sonar. sion to polyethylene is so difficult that finishes generally will not stay on the fibers. All Spectra fibers are supplied with ’low percent process 10.3.5 TREATMENTS finishes’ whose purpose is simply to aid in Because of their chemical inertness, polyethyl- holding the filaments together in the bundle. ene fibers bond poorly to matrices, with Like aramid fibers, polyethylene fibers may consequent negative effects upon the mechan- be twisted, particularly for applications such as ical properties of their composites. Surface marine cables. For bare fiber, twist initially treatment by acid etch, plasma etch or corona increases the strength, although it decreases discharge can significantly improve the the modulus of the fiber. For Spectra fiber, the Table 10.11 Availability of commercial polyethylene fibers
Product (reference)
Count
Hoechst Celanese fiber (Adams, 1993)
yarn pm (IO”in)
444 (400) 888 (800) 1780 (1600) 100 200
(90)b (180)
Spectra 900 (Allied, 1990, 1993)
722 (650) 1333 (1200) 5333 (4800)
60 120 480
38 (1.50) 38 (1.50) 38 (1.50)
Spectra 1000 (Allied, 1990, 1993)
239 (215) 417 (375) 722 (650) 1444 (1300)
60 60 120 240
23 30 28 28
5.6 (5) 22.2 (20) 111 (100)
1 1 1
555 (500) 1110 (1000)
50 100
Tekmilon (Mitsui, 1989) Monofilament
Multifilament a
Cornmen t s
number/ diameter dtex (den)
Dyneema SK60 (Dyneema, 1987)
Fihmen t
All Spectra fibers are supplied untwisted and with only sufficient finish to hold the fiber bundle together
(0.91) (1.18) (1.10) (1.10)
27 (1.06) 54 (2.13) 121 (4.76) =38 ( ~ 1 . 5 ) =38 (~1.5)
All availabilities are subject to market conditions and should be verified with the manufacturer. This fiber is newly commercially available. It will be supplied in multiples of 100 dtex.
232 Organicfibers strength of the bare fiber is optimized at a twist multiplier of 3, but at this value there is also about a 20% loss in modulus. As with aramids, however, twist is not desirable in fibers used in composites. Polyethylene fiber manufacturers provide most of their fibers untwisted. 10.3.6 FORMS AND AVAILABILITY
Table 10.11 lists the availability of most of the commercial high performance polyethylene fibers. In most cases, each denier of fiber is available in only one package size and weight. The mechanical properties of different deniers may vary somewhat from the nominal values given in Table 10.9. In addition to the fiber forms listed in Table 10.11, Dyneema SK60 is available in various fabrics. Spectra fibers are also available as chopped fiber in lengths from 6 to 20 rnm (0.25 to 0.8 in) and in fabrics of various weaves. Tekmilon is available as monofilament, multifilament and tape. Information about the availability of the fibers shown in Table 10.11; information about other products and about special formulations can be obtained from the sources given in Table 10.12.
Pricing Spectra fibers are currently priced from about $15/lb for larger denier fibers to about $45/lb for the small denier fibers. Dyneema fibers are not marketed in the USA. Because of the strong yen, at the time of writing Tekmilon fibers are significantly more expensive than Spectra fibers in the USA. The new melt-extruded polyethylene fiber from Hoechst Celanese is designed to offer good properties at a lower cost than the gel spun fibers. It costs 15-50% less than the comparable Spectra fibers. 10.3.7 DESIGN CONSIDERATIONS
Like aramid fibers, polyethylene fibers are mainly used in applications where they offer a unique combination of properties. This property combination includes outstanding specific strengths and stiffnesses, high toughness, outstanding abrasion resistance and very low density. In any design, the relatively low melting point and low use temperature of polyethylene fibers as well as the relatively poor creep resistance must be considered. The creep resistance is improved in a fiber such as
Table 10.12 Sources of information on commercial polyethylene fibers Product
Information source
Dyneema fibers
Dyneema Vof, Dr. Nolenslaan 119A, PO Box 599,6130 AN Sittard, The Netherlands Dyneema Japan Ltd., 2-8, Dojima Hama 2-chome, Kita-ku, Osaka 530, Japan
Hoechst Celanese’s high performance polyethylene fiber
Hoechst Celanese Corporation, PO Box 32414, Charlotte, NC 28232-2414, USA
Spectra fibers
Allied Fibers, Allied-Signal Inc., High Performance Fibers Technical Center, PO Box 31, Petersburg, VA 23804, USA Mitsui Petrochemical Industries Ltd., Advanced Materials and Products Department, Kasumigaseki Bldg., 2-5, Kasumigaseki 3-chome, Chiyodaku, Tokyo 100, Japan Mitsui Petrochemicals (America), Ltd., 1000 Louisiana, Suite 5690, Houston, TX 77002, USA
Tekmilon fibers
Extended chain polyethylene fibers 233 Allied-Signal’s Spectra 1000, which has significantly improved creep resistance compared to Spectra 900. Polyethylene fibers are often hybridized with other more creep-resistant fibers in applications where prolonged loading is anticipated. Without hybridization polyethylene fibers must be limited to applications where long term, high load level, or elevated temperature loading is not anticipated, or where creep is otherwise not of concern. As with aramids, the poor off-axis and compressive properties of polyethylene fiber composites may also be concern. This is not a problem, of course, in applications where the loading is mainly in axial tension. The X-ray and radar transparency of the fiber can present significant design advantages. The UV resistance of the fiber is very good and thus does not present a design problem in most applications. Because of its chemical inertness, polyethylene fibers are almost impossible to dye, although color can be added during the fiber spinning process. As with aramids, polyethylene fibers are optically opaque, so resin penetration within a composite cannot be determined visually. Polyethylene fibers can be used with a variety of resins, including polyurethanes, epoxies, vinyl esters, polyesters and thermoplastics, so long as the composite can be processed below 120°C (250°F).Polyesters are economical, vinylester resin systems provide outstanding impact properties and epoxies give better translation of structural properties. The preferred thermosetting matrix cure temperature is 93-104°C (200-220°F) (Allied, 1990). The fiber manufacturers can be very helpful in choosing an appropriate resin system for an application. As with aramids, polyethylene fiber composites are relatively hard to machine and producing a smooth final machined surface requires special techniques. Machining techniques developed for aramid composites can be used successfully, as can hot knife or hot wire cutting (Allied, 1989).
Polyethylene fibers can be damaged by rough handling and should be handled with care before and during processing. They present minimal safety or environmental concerns and most are biocompatible, offering another potential design advantage. 10.3.8 APPLICATIONS
Most of the current applications of polyethylene fibers are not in structural composites. One of the main uses is in ropes and cables, particularly in marine and off-shore applications. The fibers are used because of their high strength, outstanding abrasion resistance (up to ten times that of aramids), low density (since they float), good UV stability, resistance to seawater and high durability. The fibers are also used as marine sewing threads. In another major application, UV- and water-resistance are again important. Both Spectra and Dyneema fabrics are used, typically with a film coating, in sails. Unlike aramids, polyethylene sail can be folded and repacked numerous times without damage. This latter quality also led to the selection of Spectra fabric for the anchor balloon of the Hilton Earthwinds round-the-world balloon flight project and to their usage in lightweight, durable backpacks. Polyethylene fabrics are used as filter cloths, where the excellent chemical resistance is a tremendous advantage. Spectra fabric has been used in oil containment and recovery systems following the Persian Gulf war. The fabric is treated so that water passes through it but oils and other floating pollution do not. Because of their excellent biocompatibility, polyethylene fibers are used as sutures and as artificial ligaments. They are also used in surgical gloves, because of their excellent cut resistance, biocompatibility and low absorption of fluids. While the poor temperature resistance rules out the use of polyethylene fibers in thermal protection, they are used in industrial protective clothing and in ballistic protection and impact shields, with or without
234 Organic fibers
a matrix. Allied makes a special non-woven polyethylene fabric called Spectra Shield@, which has alternating unidirectional layers held together by a polymer matrix and gives outstanding performance in such applications. As composites, polyethylene fibers and fabrics are used in boat hulls, water skis, sailboards, canoes and kayaks. They have been explored for sporting goods ranging from archery bows to ski poles. In all these examples the excellent impact resistance they impart to their composites is another significant advantage. In applications such as skis and tennis rackets, as well as speaker cones, the excellent vibrational damping capability is also an advantage. However, in spite of this superior performance, the much higher cost of polyethylene compared to glass has limited the usage in boat hulls and sporting goods to high end applications such as racing competitions where performance and/or safety are more important than cost. In some cases, however, the addition of polyethylene fiber actually lowers the overall cost of the product. One such application is in wrapping ice hockey sticks with Spectra fiber to improve their durability. In this case, the additional cost to wrap the stick is more than offset by the longer life achieved. Polyethylene fibers are also used to reinforce rubbers and elastomers. In many of these applications, their excellent vibrational damping characteristics are important. However, the temperature sensitivity limits the usage in tires to off-road vehicles (Dyneema, 1987). Because of their X-ray, radar and sonar transparency and low dielectric constant, they have significant potential in applications such as radomes, sonar domes and X-ray tables. As mentioned before, polyethylene can be hybridized with other fibers to provide significant improvements in impact resistance. Current hybrid applications range from bike frames to impact shields.
CONCLUSIONS
For many applications extended-chain polyethylene fibers are superior to all other fibers, particularly when properties such as toughness, dielectric constant, and/or hydrolytic stability are of concern. However, because of their relatively low melting temperature, polyethylene fibers must be limited to moderate temperatures and to applications where the creep response is acceptable. For sail-cloth, marine rope, pressure vessels and other applications where the service temperature is not a governing factor, polyethylene fibers can be expected to make serious inroads into or even dominate their respective industries. However, they cannot supplant aramid, carbon, or glass fibers in applications where elevated service temperature or creep resistance are critical. 10.4 OTHER ORGANIC FIBERS 10.4.1 AROMATIC POLYESTER FIBERS
Aromatic polyester fibers are prepared by spinning from a liquid crystalline melt followed by heat treatment to form high strength, high modulus fibers. While many such fibers have been synthesized since the late 1970s, the only fiber commercially available in the USA today is Vectran@fiber from Hoechst Celanese Inc. The general structural formula of this fiber is shown in Fig. 10.12. Vectran was developed in the 1970s in response to tire customers who wanted equivalent performance to aramid fibers but at a lower
p: Y
Fig. 10.12 General structural formula of Vectran. aromatic polyester fiber.
Other organic fibers
cost. Hoechst Celanese was successful in providing a fiber with the desired performance, but unfortunately the resulting cost was even hgher than the aramids. Shortly thereafter, Hoechst Celanese stopped marketing Vectran fiber and marketed Vectran resin instead. Vectran resin quickly became the material of choice in the electronics and computer industries for small, very close tolerance connectors, plugs and other components. Based upon this success, in 1989 Hoechst Celanese reintroduced Vectran as a fiber product. Since the fiber is more expensive, by 1.5to 3 times, than aramids, the marketing focus is on areas where aramids do not meet the performance requirements (Adams and Farrow, 1993a). Vectran is a polyester-polyarylate fiber. Unlike the aramids, Vectran melts at high temperature. It is melt spun on conventional polyester spinning equipment and the as-spun fibers are then heat treated in a sequence of steps (Adams and Farrow, 1993b).It is the only commercially available melt-spun liquid crystalline polymer fiber (Hoechst Celanese, 1990).
Properties Vectran HS offers a unique combination of properties: high strength, no creep, low moisture absorption, negative coefficient of thermal expansion, good property retention
235
over a broad temperature range and excellent chemical resistance. The density of Vectran HS is 1.41 g ~ m - ~ (0.0509 lbs in") (Hoechst Celanese, 1990). It melts at 330°C (636°F). Like aramid fibers, Vectran HS has a negative axial coefficient of thermal expansion. From 20°C (68°F) to 145°C (293°F) its longitudinal linear thermal expansion coefficient is 4 . 8 x lo4 /"C ( - 2 . 7 ~ /OF). The coefficient increases to -14.6 x 10" /"C (-8.1 x lo4 / O F ) from 145°C (293°F) to 200°C (392°F) and to -26.7 x lo4 /"C (-14.8 x / O F ) from 200°C (392°F) to 290°C (554°F) (Beers and Ramirez, 1990).It has good temperature resistance, although not as good as aramid fibers since it melts at high temperature. Its shrinkage in hot air at 177°C (350°F)or in boiling water is less than 0.5%. Vectran HS fiber is outstanding in its mechanical properties. Its axial mechanical properties are summarized in Table 10.13. Vectran displays no creep when tested for 2760 h at 50% of its ultimate tensile strength. This behavior is significantly better than both aramid and polyethylene fibers. Vectran also has excellent vibrational damping characteristics, better than aramids. Vectran HS has superior abrasion resistance to Kevlar 29, although not as good as polyethylene fibers. In flexural fatigue, Vectran HS braid exhibited a 10% reduction in strength after one million
Table 10.13 Axial tensile mechanical properties of representative non-aramid, non-polyethylene organic fibers Fiber
Reference
Specific gravity
Tensile modulus GPa (Msi)
Tensile strength, MPa (ksi)
Elongation at break, %
Vectran HS
Hoechst Celanese, 1990
1.41
64.8
2840 (412)
3.3
(9.4)
Adams and Farrow, 1993a
1.41
62-86 (9.0-1 2.5)
2500-3100 (363450)
2.2-2.5
PBO (Dow)
Burk, 1993
1.56
152 (22)
5650 (820)
3.5
PBO, high modulus (Dow)
Burk, 1993
1.56
276 (40)
5520 (800)
1.5
236 Organicfibers cycles (and maintained this strength level to five million cycles), while a Kevlar 29 braid showed a 30% strength reduction under the same conditions (Beers and Ramirez, 1990). Vectran absorbs very little water, having a moisture regain of less than 0.1%. It is hydrolytically stable. It has excellent chemical resistance being resistant to organic solvents, to acids at less than 90% concentration and to bases at less than 30% concentration (Hoechst Celanese, 1990). Vectran's dielectric constant is 3.3at 1 kHz (Hoechst Celanese, 1990). Like aramid and polyethylene fibers, Vectran is difficult to cut and its composites are difficult to machine. Typical aramid composite machining techniques can be used successfully. Forms, availability and treatments Vectran is produced in the USA by Hoechst Celanese Corporation and in Japan by the Kuraray Corporation under license from Hoechst Celanese. It is available as Vectran HS, a high strength reinforcement fiber, Vectran M, a high performance matrix fiber, and as engineered, commingled combinations of Vectran HS with Vectran M and S-2 Glass fiber with Vectran M. Vectran HS fiber is available in dtex (deniers) of 222 (200),833 (750), 1000 (900)and 1667 (1500).These fibers are composed of 40,150,180 and 300 filaments respectively, with the filaments being 23 pm (0.91x 10" in) in diameter. Vectran HS fibers are offered with or without a standard textile finish to assist in processing and/or to provide (dramatically improved) abrasion resistance. Commingled Vectran HS or M fibers have no finish. (In the commingled S-2
glasslvectran M products, the glass fibers are producer-sized.) (Hoechst Celanese, 1990). More information on Vectran can be obtained from the source listed in Table 10.14. Applications As with aramid and polyethylene fibers, ropes and cables are an important usage. Marine cables, fish nets, towing ropes, cargo tie downs, slings, sails, bicycle brake cables and optical fiber reinforcement all have been made from Vectran HS. Olympic target archers use bow strings from Vectran HS having a proprietary abrasion resistant finish. The result is increased arrow speed with no creep of the string. At the last America's Cup yacht races, Vectran HS was used in at least six yachts, either in sails or in marine cables. Vectran sails stretch far less than either aramid or polyethylene and they have four to six times the life of aramid sails (Adams and Farrow, 1993a). Safety materials and protective garments have also been made of Vectran in industries ranging from meat packing to metal working. In these applications, Vectran is superior to aramids, which have poor resistance to bleach, and to polyethylene fibers, which are sensitive to the high temperatures used in drying laundered garments. Vectran composites have been used in aerospace applications and in recreation and leisure applications such as canoes, golf clubs, baseball bats, hockey sticks, tennis rackets, bicycles, skis, ping pong paddles, paragliders and stereo speaker cones. Vectran properties of importance in these applications include the low moisture absorption, the excellent damping characteristics, high stiffness, lack of creep
Table 10.14 Sources of information on non-aramid, non-polyethylene organic fibers Product
Itzj?ormation source
Vectran fibers
Hoechst Celanese Corporation, P.O. Box 32414, Charlotte, NC 28232-2414, USA
PBI fibers
Hoechst Celanese Corporation, P.O. Box 32414, Charlotte, NC 28232-2414, USA
PBO fibers
The Dow Chemical Company, Midland, MI 48674, USA
Other organic fibers 237 and good flexural fatigue properties. Vectran is used where the cost of the fiber is secondary to its performance (Adams and Farrow, 1993a). Hybrid tennis rackets have been made by Prince Manufacturing Company and Dunlop. Vectran HS is combined with carbon fiber to give greater speed and power with vibration characteristics as good as wood. Jennifer Capriatti played with a Vectran racquet at Wimbledon in 1992. Other actual or potential uses include antenna guy wires, chemical resistant packings and gaskets, heat and creep resistant belting, medical and surgical equipment, pressure vessels, printed circuit board substrates and aerial tow ropes. 10.4.2 AROMATIC HETEROCYCLIC POLYMER FIBERS
Two aromatic heterocyclic polymer fibers are currently available or in development in the United States. These are PBI fiber from Hoechst Celanese and PBO fiber from Dow Chemical Company.
Fig. 10.14 Structural formula of poly-p-phenylene benzobisoxazole (PBO) (Yang, 1992).
in 1983. It has excellent chemical and solvent resistance and does not burn. It is, however, more expensive than the aramids and has an intrinsically high moisture absorption. It is used mainly in woven form in fireblocking layers, including aircraft seat cushions and fire-fighting overgear. It was also used in chemical warfare suits in Operation Desert Storm. In order to reduce cost, PBI is also used in blends with aramids for thermal protective apparel. PBI has potential applications as a fiber reinforcement in composites, but currently its only composites application is as a matrix resin or as a matrix-precursor for carbon-carbon composites (Yang, 1992; Conrad, 1993). More information on PBI can be obtained from the source given in Table 10.14.
PBI fiber
PBI fiber is produced from a high performance polybenzimidazole. Chemically it is poly-2,2'rn-phenylene-5,5'-benzimidazole,with the structural formula shown in Fig. 10.13. The fiber was commercialized by Hoechst Celanese
PBO fiber
PBO fiber is a polybenzoxazole, specifically poly-p-phenylene benzobisoxazole, with the structural formula shown in Fig. 10.14. PBO fiber resulted from a US Air Force program aimed at developing high strength fibers for advanced composites. In the late 1980s, Dow Chemical purchased worldwide rights to the polymer. Dow has now constructed pilot plant facilities for monomer, polymer and fiber and the fiber is available for evaluation in pre-production quantities (Burk, 1993). As with aramids, PBO fibers are spun from a liquid crystalline solution using dry-jet wet spinning. This is, however, a more difficult process than for aramids. The fiber is then heat Fig. 10.13 Structural formula of poly-2,2'-rn-pheny- stretched to improve its orientation and properties (Wolfe, 1990). lene-5,5'- benzimidazole (PBI) (Yang, 1992).
238 Organic fibers PBO is one of the most thermally and thermo-oxidatively stable organic polymers known. No weight loss was observed for PBO held at 316°C (600°F) (Wolfe, 1990)and weight loss of only O.O6%/h was observed at 370°C (700°F) (Burk, 1993). Its decomposition temperature is 600°C (1110°F) (Burk, 1993). Exposed to flame, PBO chars, but does not support combustion. (Wolfe, 1990) Dow's PBO fiber has a longitudinal coefficient of thermal expansion of -6 x 104/OC (-3.3 x lO"/"F) (Burk, 1993). PBO fiber has a significantly higher tensile strength and modulus than any other known organic fiber. PBO fibers have been produced with tensile moduli of as high as 470 GPa (68 Msi). Dow's current pre-production fibers do not achieve these high levels, but do nonetheless have excellent axial mechanical properties, as shown in Table 10.13. However, like all other high performance organic fibers, PBO fibers are quite weak in compression, with a fiber compressive strength comparable to that of aramids (Burk, 1993).They also bond poorly to epoxy matrices, so their off-axis properties are also poor (Wolfe, 1990). For these reasons, as with other organic fiber composites, PBO composites are limited to applications where structural loading is mainly in axial tension. Moisture regain for Dow's PBO is 2.0% for the standard fiber and less than 0.570 for the high modulus version. The moisture resistance is significantly better than aramids (Burk, 1993).PBO is highly resistant to hydrolysis, acid chemical attack, bases, solvents, electron bombardment and laser radiation. Its UV stability is outstanding (Wolfe, 1990). Dow's PBO fiber has a lower and more stable dielectric constant than that of aramids, 3.0 at 100 kHz (Burk, 1993). The price for commercial PBO will be volume dependent, but will be higher than that for aramids. PBO fiber will be used where aramids and other fibers do not meet the performance needs, particularly for strength, modulus and flammability. Potential applica-
tions include composites loaded in tension, such as pressure vessels, missile cases and tensile beams. PBO fiber composites may also be used in non-load-bearing applications where high temperature exposure or harsh chemical environments are anticipated, such as rocket insulation systems and brake and transmission systems. The high strength could also lower the weight of composites used in spacecraft and in recreation and sporting goods. PBO also has significant potential application to ballistics, where, as a fabric or composite, it performs equally well at half the weight of an aramid. PBO composites could provide outstanding containment systems for high speed rotors and turbines where high temperature exposure is of concern. The fibers also have potential for bomb containment systems, for fire resistant and cut resistant apparel and fire blocks, as well as ropes and cables (Burk, 1993). More information on PBO can be obtained from the source given in Table 10.14. 10.5 CONCLUSIONS
While high performance organic fibers are not competitive with inorganic fibers in all of their properties, they offer certain properties and combinations of properties that are unavailable with inorganic fibers. All suffer from certain limitations, such as poor off-axis and compression properties and/or temperature limitations. However, if these limitations are properly considered, high performance organic fibers can make possible designs that can be achieved in no other way. REFERENCES Abbott, N.J., Donovan, J.G., Schoppee, M.M. and Skelton, J. 1975. Some mechanical properties of Kevlar and other heat resistant, nonflammable fibers, yarns, and fabrics. Technical Report AFML-TR-74-65, Part 111. Wright Paterson Air Force Base: Air Force Materials Laboratory. Adams, P.M. 1993. Private communication. 10-28-93. Charlotte, NC: Hoechst Celanese Corporation.
References Adams, P.M. and Farrow, G. 1993a.Advanced fiber materials for specialty applications from fully aromatic polyesters. Presented at Textile Research Institute, 63rd Annual Conference, 5-6 May 1993, Princeton, NJ. Adams, P.M. and Farrow, G. 1993b. Processing, properties and applications of fibers from fully aromatic polyesters. Unpublished paper. Charlotte, NC: Hoechst Celonese Corp. Adams, P.M. Farrow, G. and Beers, D. 1995. Advanced fiber applications: properties and applications of fibers from fully aromatic polymers. TAPPI J., 78(11), 169-174. Adams, W. Wade and Eby, R.K. 1987. High-performance polymer fibers. MRS Bulletin 12 (12): 22-26. Akzo. 1990. The aramid fiber for high-performance composites Twaron. Amhem, The Netherlands: Akzo Fibers and Polymers Division. Akzo. 1991. Twaron product information yarns, fibers and pulp. Arnhem, The Netherlands: Akzo Fibers Division, Aramid Fibers. Allied. 1989 (received). Spectra high performance fibers for reinforced composites. Undated. Petersburg, VA: Allied Fibers, Allied-Signal, Inc. Allied. 1990. Spectra high performance fibers. Petersburg, VA: Allied Fibers, Allied-Signal, Inc. Allied. 1991a. Creep at 10% load (room temperature). LB006. 6/17/91. Petersburg, VA: Allied Fibers, Allied-Signal, Inc. Allied. 1991b. Creep at 30% load (room temperature). LB003. 6/91. Petersburg, VA: Allied Fibers, Allied-Signal, Inc. Allied. 1991c. High temperature creep. LB005.6/91. Petersburg, VA: Allied Fibers, Allied-Signal, Inc. Allied. 1991d. Spectra creep 20% load 5°C. LB004. 6/91. Petersburg, VA: Allied Fibers, AlliedSignal, Inc. Allied. 1991e. Extended chain polyethylene tensile properties at temperature. LB007. 7/91. Petersburg, VA: Allied Fibers, Allied-Signal, Inc. Allied. 1993 (received). Spectra high performance fibers, product specifications. Undated. Petersburg, VA: Allied Fibers, Allied-Signal, Inc. Beers, D.E. and Ramirez, J.E. 1990. Vectran highperformance fibre. J. Textile Institute 81 (4): 561-574. Blades, H. 1973. US Patent. 3 767 756. Blades, H. 1974. US Patent. 3 817 941. Burk, W.R. 1993. Private communication. 7-16-93. Midland, MI: The Dow Chemical Company. Chiao, C.C., Sherry, R.J. and Chiao, T.T. 1976. Strength retention and life of fiber composite
239
materials. Composites 7 107-109. Chiao, T.T., Chiao, C.C. and Sherry, R.J. 1976. Lifetimes of fiber composites under sustained tensile loading. UCRL- 78367. Livermore, CA: Lawrence Livermore Laboratory. Chiao, C.C. and Chiao, T.T. 1982. Aramid fibers and composites. In Handbook of Composites, ed. George Lubin, pp. 272-317. New York: Van Nostrand Reinhold. Clements, L.L. and Moore, R.L. 1977. Composite Properties of an Aramid Fiber in a RoomTemperature-Curable Epoxy Matrix. SAMPE Quarterly 9: 6-12. Conrad, D. 1993. Private communication. 7-19-93. Charlotte, NC: Hoechst Celanese. DeCos, L. 1993. Private communication. 9-27-93. Conyers, GA: Akzo Fibers, Inc. DuPont. 1981. Properties of Nomex aramid filament yarns. Bulletin NX-17. December 1981. Wilmington, DE: E.I. DuPont de Nemours & Co. Inc. DuPont. 1983. Kevlar aramid, the fiber of choice in boat hull reinforcement. E-46814. 10/83. Wilmington, DE: E.I. DuPont de Nemours & Co. Inc. DuPont. 1986. Data manual for Kevlar 49 aramid. May 1986. Wilmington, DE: E.I. DuPont de Nemours & Co., Inc. DuPont. 1989 (received). Kevlar aramid, the uncommon material for uncommon solutions. H-05500-1. Undated. Wilmington, DE: DuPont, Fibers Department. DuPont. 1990. Presenting Kevlar 119 aramid fiber, for longer service life in demanding applications, 5/90. Wilmington, DE: DuPont Fibers Department. DuPont. 1992a. Kevlar aramid fiber technical guide. 12/92. Wilmington, DE: DuPont Fibers, Kevlar Products. DuPont. 199213. Kevlar aramid, properties and uses of Kevlar 29 aramid, Kevlar 49 aramid, Kevlar 68 aramid in fiber optic and electromechanical cables. Information bulletin K- 506C, revised November 1992. H-37390. Wilmington, DE: DuPont Fibers Department. DuPont. 1992c. Internal price list, Kevlar yarn, 11/23/92. Wilmington, DE: DuPont. DuPont. 1992d (received). Kevlar aramid KM2, preliminary information bulletin, H-35645. Undated. Wilmington, DE: DuPont Fibers, Kevlar Products. DuPont. 1993a. Prices for Kevlar 29 yarns used in textile processing. Price list effective 1/4/93.
240 Organic fibers Wilmington, DE: DuPont. DuPont. 199313. Prices for Kevlar 29 yarns used in ropes and cables. Price list effective 1/4/93. Wilmington, DE: DuPont. DuPont. 1993c. Prices for Kevlar 49 yarns used in textile processing. Price list effective 1/4/93. Wilmington, DE: DuPont. DuPont. 1993d. Prices for Kevlar 49 yams used in fiber optics. Price list effective 1/4/93. Wilmington, DE: DuPont. DuPont. 1993e. Prices for Kevlar 49 yarns used in ropes and cables. Price list effective 1/4/93. Wilmington, DE: DuPont. DuPont. 1993f. Rope and cordage products. H37399. 2/93. Wilmington, DE: DuPont Fibers Department. DuPont. DuPont. 1993g. Properties of DuPont industrial filament yarns, DuPont nylon, Cordura nylon, Dacron polyester, Nomex aramid, Teflon fluorocarbon, Kevlar aramid. Technical Information, Multifiber Bulletin X-273, April 1993. A-90240. Wilmington, DE: DuPont, Fibers Department. DuPont. DuPont. 199317. Properties and processing of DuPont Kevlar Aramid Yarn for Mechanical Rubber Goods. Technical Information, Kevlar Bulletin K-10, June 1993. Wilmington, DE: DuPont Fibers Department. DuPont. 1993i (received). The second generation of ballistic protection, new Kevlar 129. H-13653. Undated. Wilmington, DE: DuPont. Dyneema. 1987. Dyneema SK60, high strength/high modulus fiber, properties and applications. Sittard, The Netherlands: Dyneema Vof. Gerzeski, Roger H. 1989.Vniivlon/polyamidobenzimidazole - USSR’s aramid fiber forming polymer. In Reference Book for Composites Technology, Vol. 1, ed. Stuart M. Lee, pp. 271-325. Lancaster, PA: Technomic Publishing Company. Hoechst Celanese. 1990. Vectran liquid crystalline polymer fiber. Charlotte, NC: Hoechst Celanese Corporation. Hoechst Celanese. 1993 (received). Typical properties of high modulus polyethylene. Charlotte, NC: Hoechst Celanese Corporation. Hongu, Tatsuya and Phillips, Glyn 0. 1990. New Fibers,. London: Ellis Honvood Ltd. Jaffe, M. 1989. High-modulus high-strength organic fibers. In Concise Encyclopedia of Composite Materials, ed. Anthony Kelly, pp. 129-134. Oxford : Pergamon Press.
Kaiser VIAM. 1993a. Certificate, synthetic aramid filament, 58.8 tex. San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993b. Certificate, fiber SVM 6.3(40)Al. San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993c. Certificate, fiber SVM 14.3-A1 ‘acidic.’ San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993d. Certificate, fiber SVM 29.4(200)-Al ’ N (neutral). San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993e. Certificate, fiber SVM-K 29.4 ’acidic.’San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993f. Certificate, fiber SVM 29.4-A1 ’acidic.’San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993g. Certificate, fiber TOW SVM 58.8(300)X17- 1000; VTV heat treatment. San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993h. Certificate, fiber TOW SVM 58.8(300)X17- 1000; TOSN heat treatment. San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993i. Certificate, fiber SVM 58.8 ’acidic.’ San Leandro, CA: Kaiser VIAM. Kaiser VIAM. 1993j. Certificate, fiber SVM 6.3(40)Al. San Leandro, CA: Kaiser VIAM. Kwolek, S.L. 1971. U.S. Patent 3 600 360. Kwolek, S.L. 1972. US Patent 3 671 542. Kwolek, S.L. 1974. US Patent 3 819 587. Mahn, Harry M. 1993. Director R & D. Private communication. 9-27-93. New York: Teijin America, Inc. Mitsui. 1989. High-performance fiber material Tekmilon. 7/89. Tokoyo, Japan: Mitsui Petrochemical Industries, Ltd. Penn, Lynn and Larsen, Fred. 1979. Physiochemical properties of Kevlar 49 fiber. J. Appl. Polym. Sci. 23: 59-73. Perkins, G. 1993. Private communication. Conyers, GA: Akzo Fibers, Inc. Prevorsek, D.C. 1989. Ultrahigh modulus/strength polyethylene fibers and composites. In Reference Book for Composites Technology, Vol. 1, ed. Stuart M. Lee, pp. 167-174. Lancaster, PA: Technomic Publishing Company. Teijin. 1989. High tenacity aramid fibre Technora. Technical Information TIE-05/89.11. Osaka, Japan: Teijin Limited. Teijin. 1991. Heat resistant aramid fiber Teijinconexa. Technical Information CN02/91.2. Osaka, Japan: Teijin Limited. Teijin. 1993 (received). Technora - a para aramid copolymer fiber. Bulletin No. L 1.0. Undated. Osaka, Japan: Teijin Limited. Ward, I.M. and McIntyre. 1986. High-modulus
Appendix
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fibers. In Encyclopedia of Materials Science and Engineering, Vol. 3, ed. Michael B. Bever, pp. 213940. Cambridge, MA: MIT Press. Weedon, G.C. and Tam, T.Y. 1986. New PE fibers for high-end composites. Modern Plastics. 63(3):64-68. Wolfe, J.F. 1990. Polybenzothiazoles and polybenzoxazoles. In Encyclopedia of Polymer Science and Technology, Vol. 11, ed. J.L. Kroschwite, pp. 601-635. New York: John Wiley. Wu, E.M. 1980. Strength degradation of aramidfiber/epoxy composites. AMMRC-TR-80-19. Watertown, MA: Army Materials and Mechanics Research Center. Yang, H.H. 1992. Kevlar Aramid Fiber. New York: Wiley-Interscience. Yang, H.H. 1989. Aromatic High-Strength Fibers. New York Wiley-Interscience.
as the weight in grams of 1000 meters of the material. A related term is decitex (dtex),0.1 tex, which is often used in order to be comparable to the US quantity 'denier.' Equilibrium moisture content: moisture absorbed by a fiber after it has been dried at 50°C (122°F)for 2 h and then equilibrated at 20°C (68°F)and 55% relative humidity Strength retention: percent of room temperature strength retained following exposure to the conditions indicated Tenacity: the ultimate failure strength of a fiber per unit original area per unit weight. The most commonly used units are 'grams per denier' (gpd) and 'newtons per tex' (N/ tex).
10.6 APPENDIX
Conversion factors
Fiber size: 1 tex = 9 denier = 10" kg/m Twist: Definitions 1 tpi (turns per inch) = 39.37 t/m (turns per Denier: Term in common usage in the fiber meter) industry in the USA to describe the fineness Twist multiplier = [t/m (dtex)'/*]/3000= [tpi of a fiber or fiber bundle. The denier is (denier)'/'/ 731 defined as the weight in grams of 9000 Modulus, stress, strength, and tenacity: meters of the material. This is also known as 1kgf / mm2 = 9.806550 MPa the 'count'. Its inverse measure is the 1 ksi = 6.894757 MPa 'yield', expressed in yards per pound or 1cN/tex = 0.01 N/tex = 10 pf MPa = 1.45 pf ksi meters per kilogram. where pf is the specific gravity of the fiber Tex: Term in common usage in the fiber indus- 1 gpd = 8.826 cN/tex = 88.26 pf MPa = 12.8 pf try outside the USA to describe the fineness ksi of a fiber or fiber bundle. The tex is defined DEFINITIONS AND CONVERSION FACTORS
PARTICULATE FILLERS Harry
11
s. Kat2
11.1 INTRODUCTION
The preceding chapters of this handbook discussed fiber reinforcements. Fibers play an important role as the outstanding means for obtaining great increases in the strength and modulus of the matrix material. The focus of attention on fiber reinforcements has usually blinded potential end users of composite materials to the many beneficial uses of nonfibrous, or particulate, fillers. However, industrial experts are now becoming more aware of these benefits and there will undoubtedly be increased particulate filler usage in future polymers, ceramics and metals. Particulate fillers can provide improved materials, as compared with the unfilled matrix and can also be synergistic with a fiber reinforcement to further improve the system performance. A fiber may be described as a particle with a length-to-diameter ratio of greater than 10 to 1. A particulate filler may be described as a non fibrous solid that fits the definition given in the Compilation of ASTM Standard Definitions, ‘a solid compounding material, usually in finely divided form, which may be added in relatively high proportions to a polymer for technical or economic reasons.’ This chapter will not include additives such as antioxidants or internal mold release agents, which serve important compounding functions, but are used at low levels, usually below
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
2 wt.% in the composite material. The only exception is the inclusion of a brief discussion of coupling agents, which serve an important function in achieving optimum physical properties in particulate filler composites. There are many good reasons for using particulate fillers in plastic, metal or ceramic matrices, in addition to the obvious usual reduction in cost of the final product. In the case of plastics, the addition of fillers provides a reduction of shrinkage during the cure of a thermoset polymer system or the injection molding of a thermoplastic resin. This reduced shrinkage results in important benefits such as avoidance of the warpage or cracking that may occur, especially in the case of large molded parts. Also, the thermal conductivities of mineral fillers are usually in the order of ten times greater than the thermal conductivity of polymers. Therefore, the filled polymer has a much greater thermal conductivity than the unfilled resin. This provides an important advantage in processes such as injection molding, where the cycle time is often determined by the time to cool the part in the mold. The faster cooling rate of the filled plastic will provide cost savings due to the faster cycle time. Another advantage of the higher thermal conductivity is the faster dissipation of localized hot spots, which could cause thermal decomposition of a polymer or failure of a sensitive electronic component adjacent to the plastic. There are many types of particulate fillers1. They may be classified as mineral, natural or synthetic, inorganic or organic. The mineral fillers include calcium carbonate, clay,
Corninon particulate fillers feldspar, nepheline syenite, talc, alumina trihydrate, natural silicas and mica. The importance of particulate fillers is readily apparent:
243
11.2.1 CALCIUM CARBONATE
Calcium carbonate is one of the lowest cost particulate fillers. Price varies depending on the grade and quantity purchased. During ’In 1993, about 4.9 billion pounds of fillers and 1994, the listed price for good quality standard extenders, which included minerals, synthetic uncoated grades ranged from about $80 to and organic based materials, were used by the $200 per ton. Special grades, such as precipiUS plastics industry. The estimated value of tated and coated can cost over 25 cents per these materials was $1.2 billion. Minerals rep- pound. Calcium carbonate is non-toxic, nonresent the majority of these products. The irritating, odorless and is white in color with a annual growth rate of synthetic fillers was esti- low refractive index. It is nonabrasive and is the standard material for Mohs’ 3 rating on the mated at 7%.12 Mohs’ hardness scale. There is an abundant Manufactured fillers include products such supply of this mineral filler and it is commeras glass beads, metal powders and synthetic cially available in a wide range of particle silicas. sizes. It exhibits a low oil absorption rating There are many particulate organic fillers and is readily dispersed at high loadings in such as wood flour, carbon black, various liquid systems. As is the case with other parstarches, ground rice hulls, peanut shell and ticulate fillers, its use reduces shrinkage reclaimed rubber. during molding and curing. Because of this Low density particulate fillers are available characteristic, it is a usual filler in low shrink and these include perlite and hollow microsreinforced polyester sheet molding compheres of glass, ceramic, or organic polymers pounds. Even with high loadings, it provides such as phenolic and polyvinylidene chloride. relatively low stiffness to the composite. It is Specific particulate fillers can give a comstable up to 800°C. posite material special characteristics such as electrical conductivity, biodegradability, thermochromic, photochromic, low surface 11.2.2 FELDSPAR AND NEPHELINE Feldspar and nepheline are similar minerals friction or magnetic properties. and serve the same purposes. These are anhydrous, alkali, aluminosilicates having excellent 11.2 COMMON PARTICULATE FILLERS chemical resistance and are offered in coarse to There are many commercial particulate fillers medium particle sizes. Because the index of and the end user must take into consideration refraction of this mineral is equivalent to many facts before deciding whether to use a polyvinyl chloride and other polymers, its use particulate filler and which material will be best can result in retention of transparency or for the application. The following paragraphs translucency in most polymers. Thus, less present some information as a starting point for color pigment may be required to obtain a possible selection of one of these materials for a given intensity. These mineral particulates specific end use. Most of these descriptions are usually provide ease of wetting and are readcondensationsof the more complete details that ily dispersed and de-aired. They cause little to can be found in The Handbook of Fillers for no effect on promoters, accelerators, or addiI’lastics’. It should be noted that the descrip- tives that are used in liquid polymer systems. Because of the low oil absorption values, tions are mainly aimed at the use of these materials in a polymer matrix, but some of high loadings of these minerals can be added these particulate fillers may have similar uses in to liquid polymer systems without running into the problem of excessively high viscosity. metal or ceramic matrix composites.
244 Particulatefillers Feldspar and nepheline syenite are not considered to be a health hazard or toxic and these minerals have been recommended as suitable for use in composites that contact food. Feldspar is obtained from white granite by quarrying the rock and crushing it. Nepheline is the mineral derived from syenite rock and is referred to as nepheline syenite. High-grade deposits consisting of millions of tons are mined and processed northeast of Toronto, Canada. 11.2.3 KAOLIN (CLAY)
Kaolin is the common term for the mineral kaolinite, a hydrous aluminosilicate mineral, one of a number of minerals called clays. It is available in two primary grades: the natural hydrous form and a calcined, anhydrous form. The larger average particle sizes consist of mixtures of blocky and platy particles. The finer fractions are almost entirely thin plates. Clay is a rock term applying to soft, earthy ores which are plastic when mixed with water. The hydrous forms are non-abrasive, chemical resisting and have relatively high surface area, which results in high viscosities in liquid polymer systems. They disperse readily in most plastics, especially with the aid of dispersants or surfactants. In reinforced and thermoset plastics, they control flow properties to provide more uniform composite properties. Calcined kaolin is considerably harder and provides improved electrical properties in thermosets and thermoplastics. Many surfacetreated types of each grade are available to provide outstanding water resistance, electrical properties and reinforcement, in a variety of systems. 11.2.4 SILICA
There are over 20 distinct phases of silica and each is chemically SiO,. A number of them are useful as fillers in plastics. These different phases offer advantages and disadvantages, for example a synthetic colloidal form, or
amorphous phase, of silica has very high surface area. When added to liquid systems, this type of product can provide a thixotropic effect and reinforcing properties in cured rubbers or polymers. Flocculated varieties of the colloidal form give flatting characteristics and composite properties different from the particulate colloidal phase. The synthetic forms are so fine in particle size that they offer little in the way of hardness in the finished composites. Natural silica products, as in the case of the low quartz phase, come from quartzite and tripolitic forms, which are generally termed natural, microcrystalline quartz. These are usually less expensive than the synthetic, colloidal forms. They offer excellent electrical insulation, good mechanical properties and lower surface area, hence better rheological flow. Drawbacks are increased wear of processing equipment and the lack of reinforcing properties. Fused silica is manufactured by subjecting high purity quartzite to temperatures in excess of 2200°C (4000°F).Finely divided fused silica has a very low coefficient of thermal expansion. This is desirable in composites subjected to wide variations in temperatures. Silica surfaces can be modified to offer organofunctionality,hydrophobicity and nonfunctionality. commercial activity in surface modification is limited now, but the future will surely see more functional particulate silica materials with the interface tailored to specific end uses. 11.2.5 ANTIMONY OXIDE
Antimony oxide is an opaque white mineral powder which has been used as a pigment because of its good color and high hiding power. It is no longer competitive with more economical white pigments and its high cost and density makes it unattractive as an ordinary filler or extender. The major use now is as a flame retarder. Synergism between antimony oxide and halogens accounts for the widespread use of
Common particulate fillers 245 antimony oxide in plastics. Antimony oxide is added at 1-15 wt.% and affects physical properties in the same way as other fine particle size minerals. The halogen compound has a greater effect on plastic properties because larger quantities are added. The halogen compound is a plasticizer, which lowers modulus and strength. An additive system must be selected for the fabrication method to be used and for flammability and physical requirements of the specific application.
mal conductivity to a polymer composite. In each application, the grade and manufacturer must be carefully selected in order to obtain optimum performance. There are a wide variety of commercial carbon black products. Differences include particle size and particle size distribution, specific surface area, surface structure, amount of agglomeration, moisture content and minor contaminants, such as metals, sulfur, oxygen and hydrogen. These differences originate from the fact that various hydrocarbon raw materials and production methods are used to 11.2.6 ALUMINA TRIHYDRATE produce different grades of products, which Alumina trihydrate (Gibbsite)is a refined min- include furnace, thermal, acetylene, lampblack eral which offers a bonus of flame retardance and channel blacks. Carbon blacks exhibit pH and smoke suppression in most plastics appli- of 2-9 when slurried in water. cations. It is available in different grades, which Many polymers are susceptible to phocontain various amounts of minor impurities todegradation upon exposure to sunlight or and range in particle size from coarse materials, UV radiation. An effective method for reducalmost entirely +325 mesh, to very fine prod- ing or eliminating this effect is by adding about 3 wt.% of an appropriate carbon black to ucts less than 1 Fm in mean particle diameter. Alumina trihydrate is non-toxic and rela- the polymer. There are a number of specific grades of cartively inert chemically. In electrical applications, it provides the benefit of arc and bon black, especially acetylene blacks or track resistance. Since its surface has a low furnace-produced carbons, that are very effecsorptive capacity, trihydrate has low resin tive in providing high electrical conductivity demand in liquid polyester, epoxy and acrylic in a polymer composites. High electrical consystems. Because it undergoes thermal decom- ductivity can also be obtained by use of the position, it is not suitable for processing at many different grades of graphite flakes and temperatures above about 200°C (428°F). The powders that are commercially available. endothermic evolution of water at the decomposition temperature is the basis for the fire 11.2.8 MICA retardant property of this particulate filler. Mica is a generic term for a family of potassium alumininosilicate minerals, which 11.2.7 CARBON BLACK AND GRAPHITE include muscovite, phlogopite, biotite and fluParticulate carbon is available commercially in orphlogopite. The exact composition of these many forms and from many suppliers. The minerals can vary from mine to mine. In all filler known as carbon black is usually cases, the mineral is readily ground and split obtained by the pyrolysis of hydrocarbon into thin flexible sheets or flakes. Mica flake gases or oils. Its largest end use is for filling fillers have had a high growth rate as a plastics and reinforcement of rubber polymers in filler because they can provide a much greater products such as automotive tires. However, it increase of matrix modulus than most other also finds widespread applications as a filler mineral fillers. This improvement is further and coloring agent in plastics, inks, paints and optimized when the mica surface is treated paper. It can also provide electrical and ther- with appropriate coupling agent^.^
246 Particulate fillers 11.2.9 TALC
The talc of major interest to plastics users is a finely ground product consisting of thin platelets and is white in color. Due to the platy nature of this form of talc, it is considered to be a reinforcing filler in many plastic applications. Its low cost qualifies it as an extender while its platiness qualifies it as a reinforcing filler. White, platy, fine, ground talcs are available from a number of different suppliers. Composites with platy talc exhibit higher stiffness and creep resistance at ambient and elevated temperatures, when compared with equivalent composites filled with a particulate filler such as calcium carbonate. For example: polypropylene filled with a 40% loading of talc, the flexural modulus or stiffness triples from about 1380 MPa (200 000 psi) for the unfilled polypropylene to about 4140 MPa (600 000 psi) for the talc-filled equivalent. A 40"/0 loading of calcium carbonate only increases the modulus to about 2760 MPa (400 000 psi). High loadings of talc are usually accompanied by a reduction of the impact strength of the polymer matrix composite. However, this can be minimized through proper selection of the particle size and size distribution, surface treatments on the talc and resin formulation.
ture initiation of failure of a polymer matrix composite. There are many types of solid microspherical fillers, including glass, ceramics and plastics. Commercial 'A' glass spheres are available in a range of diameters, usually from 30-750 pm. Incorporating these products into a resin matrix may result in poor bonding between the resin and the sphere. For optimum composite properties, a surface treatment that is specifically tailored for the resin should be applied to the sphere. Various surface treated products are commercially available. There are many types of hollow microspherical fillers. These include glass, ceramics, polvinylidene chloride and phenolic plastic. These fillers can be used to prepare polymer composites with low densities and relatively high compressive strengths. 11.3 METHODS OF PRODUCTION
There are many different methods for production of particulate fillers. As indicated above, the mineral fillers are simply produced by mining, grinding and then removing undesirable components/contaminants and separation into various size fractions. In many cases, the end user has a choice between the natural mineral filler and a synthetic product that can provide superior 11.2.10 MICROSPHERES performance for specific applications. Calcium There are two main reasons why microsphere carbonate is the most widely used filler or extenfillers are an excellent choice for many appli- der pigment in the plastics industry. For many cations. The spherical shape provides the applications, the filler is manufactured by lowest possible surface to volume ratio, which merely crushing the mineral, followed by pullimits the area of resin adsorption. This per- verizing and air classification into the desired mits a relatively high loading of the spheres in particle size. However, a precipitated calcium a fluid matrix without the resultant high vis- carbonate is usually specified where purity, unicosity that would occur if the same volume of form small particle size and superior whiteness equivalent size non-spherical particulate were is required. The two types of calcium carbonate added. Second, the spherical shape provides a have their advantages and disadvantages. uniform stress distribution on its surface, in Different manufacturing methods will procontrast to the sharp edges of most other par- duce different size distributions of the ticulates. Sharp edges create a localized high particulate fillers and size distribution is an stress concentration, which can lead to prema- important factor to evaluate in the choice of a
Methods of production 247 filler. Manufacturers of particulate fillers present this type of data as graphs (Fig. 11.1). Each of the four calcium carbonate fillers, Atomite, duramite, Snowflake P.E. and Supermite, has quite different particle size distributions even though they are all produced by a wet grinding process. The mean particle size of these products is 3, 11,5.5 and 1 in the order cited above and the oil absorption values, by the rubout method, are 15,8,9 and 19. Note that the mean particle size of Snowflake P.E. is one half that of Duramite, whereas the oil absorption number is about the same. The relatively flat particle size distribution curve of Snowflake P.E. gives it good packing characteristics so that it provides high loading levels with low viscosity in liquid resins. Therefore, it is a good choice for a wide variety of polyester compounding methods, such as bulk molding compounds, sheet molding compounds, transfer molding compounds, pultrusion and spray up applications.
Precipitated CaCO, is produced by heating CaCO, to form CO, and CaO. The latter product is reacted with water to form calcium hydroxide, which is then recombined with the CO, to produce CaCO,. Median particle sizes are usually 0.2-2.0 pm. This material is available in high purity grades for USP, FDA and food contact applications. A disadvantage is that the cost is higher than the ground mineral products. Fumed silica is prepared by the hydrolysis of silicon tetrachloride in a flame of hydrogen and oxygen at a temperature of 1000°C or higher. While still molten, the primary small particles of silica fuse into secondary particles called aggregates. Particle size and surface area of the resultant fumed silica are controlled by varying the ratio of the reactants. After passing through a coagulation and cooling zone, the silica particles are separated from the HC1-containing combustion gases by means of cyclones or filters. The residual HC1
90 80 -
100
70
-
60
-
50
-
40
-
20 10 30
I
I
50"
I
I
I
20 Equivalent Spherical Diameter pm
Fig. 11.1 Size distributions of calcium carbonate fillers (a) Duramite, (b) Snowflake P.E., (c) Atomite, (d) Supermite. (Courtesy of ECC International, Atlanta, GA.)
248 Particulatefillers
is removed by treating the silica with moist hot air. Microsphere fillers are produced by a number of different methods. These include the fire polishing of ground particles, atomization of molten materials and calcining of crushed particles. There are a number of patents on methods for production of glass microsphere~"~. Fly ash microspheres are commercially available. These are obtained by beneficiation of the waste fly ash product that is generated in large quantities by the many utility companies that burn powdered coal.
shape from the previous two materials. This emphasizes that the choice of a particulate filler should not be made on the basis of the generic type, but must be made after evaluation of different grades that are available from
11.4 TYPES OF PARTICULATE FILLERS
11.4.1 CHARACTERIZATION BY SHAPE
The shape of a particulate filler is an important consideration in the selection of the optimum filler for a specific end use.
I 4
Fig. 11.2 Montana talc. (Courtesy of Speciality Minerals Inc.)
Microspheres As described above, spheres have many advantages over irregularly shaped particulates when used as fillers in a polymer matrix. Therefore, when particulate fillers are being considered for use in a composite material, the selection process should include the evaluation of microsphere fillers
Flakes
Fig. 11.3 Italian talc. (Courtesy of Speciality Minerals Inc.)
Flake or platelike (lamellar) particulates can provide important benefits for composite materials. In coatings, this shape of filler can lead to a highly loaded composite that is a good permeation barrier, since any penetrant must go through a tortuous path within this I material. Polymer composites, with flake reinforcements, can exhibit planar isotropy and higher modulus than composites with an equivalent loading of an irregular shaped particulate filler. The crystal form of talc may be platelike or flakes as in Montana talc (Fig. 11.2) and Italian talc (Fig. 11.3).In contrast, the New Fig. 11.4 New York talc. (Courtesy of Speciality York talc (Fig. 11.4) is quite different in particle Minerals Inc.)
Types of particulatefillers 249 various sources and companies. There can be dramatic differences in compounding characteristics and in critical properties, such as strength Or modulus~based On the specific grade of a particulate filler, how it has been processed and whether or not there is a surface treatment. In addition to talc, other flake Or Platy Some ‘lays and graphite. Mica flakes have been discussed above’ A high performance type of this material is the high ratio grades. The ratio Of flakes is defined as the ratio of the flake average diameter to the thickness. The higher aspect ratio flakes are more effective than low aspect ratio flakes in the transfer of stress from the relatively weak polymer matrix to the stronger flake reinforcement. Optimizing the properties of mica filled polymer composites has involved a number of developments, including surface treatment with coupling agents, improvements in flake separations and orientations and reduction of flake damage during processing. The impact strength of these composites is often poor, but improvements have been obtained by combinations of mica flakes with high strength short fibers and by use of polymer impact modifiers. The predominant filler for manufacture of conductive polymer composites and inks is silver flake.
11.4.2 CHARACTERIZATION BY FUNCTION
In addition to the categories or classifications discussed above, particulate fillers may be characterized by the function they serve.
Fire retardants An important type of particulate filler is the solid fire retardants discussed above. The use of these fillers will continue to increase as a means for reducing the loss of life and property due to fires in hotels and aircraft.
Lightweight fillers Particulate fillers can reduce the specific gravity of the composite material. Among the particulates for this category are perlite and the many types of hollow microspherical fillers discussed above. Perlite is an excellent low density filler, derived from a natural mineral, which is mined, ground, expanded by a heat process and screened to various size fractions. It is a multicellular material, which has the advantage that if some of the outer wall is broken, the remaining closed cell structure prevents penetration by a liquid resin. Perlite is a low cost filler, especially when evaluated in terms of cost per unit volume.
High density fillers Irregular shaped particulates Most commercial fillers may be considered as irregular in shape, even though the basic crystal structure may be a square or block type lattice. As mentioned above, mineral particulate fillers that are obtained by simple grinding and classifying into various size fractions are essentially irregular shaped particulate fillers. These fillers, such as crystalline silica or ground quartzite, have sharp edges or points that result in high stresses within a polymer composite and can lead to initiation of premature failure of the structure.
Particulate fillers may be used to increasing the specific gravity of a polymer composite. Appropriate fillers in this category are barium sulfate, nickel and lead powders.
High hardness fillers High hardness fillers such as alumina and crystalline silica may be used to increase hardness and provide abrasion resistance for polymer matrix composites.
250 Particulate fillers
High thermal conductivity fillers Beryllium oxide powder has been used in polymer matrices because it provides high thermal conductivity and c xcellent electrical resistivity. However, this powder is highly toxic and this has severely limited its usage. Alternate choices for this end use have included magnesium oxide and aluminum oxide fillers. As noted above, mineral fillers have much higher thermal conductivity than polymers, so most of these fillers will provide a polymer matrix composite with much higher thermal conductivity than the polymer matrix.
11.4.3 CHARACTERIZATION BY SURFACE PROPERTIES
The surface characteristics of a particulate filler are important in determining the final properties of the composite. Surface treatments are well established as a means for improving the processing characteristics during the addition of particulate fillers into liquid resins and for improving the physical properties and retention of the composite’s properties upon aging. Therefore, even though this chapter does not deal with additives, their importance in achieving optimum physical properties in particulate filler composites, merits brief discussion. Electrically conductive particulate fillers In composites technology, the term couElectrically conductive particulate fillers such pling agents has been used to designate as silver flake, nickel powder and graphite chemicals that are used to treat the surface of particulates can be used to provide polymer fillers and reinforcements in order to obtain matrix coatings and composites with static optimum physical properties and for longdischarge characteristics and electrical con- term retention of physical properties. Coupling agents are chemical molecules with ductivity. dual functionality, wherein one part of the molecule will adhere to one surface, e.g. filler Magnetic particulate fillers or reinforcement, while another part of the Ferrite or ceramic magnetic powders are used molecule provides a bond to the matrix mateas fillers in polymers to produce bonded mag- rial. Thus, a bonded bridge is formed between nets. The loading of these powders ranges two different materials. Coupling agents can provide benefits as from 85-90 wt.%. Plastic magnets are available in either a flexible or rigid form. processing aids for polymer matrix composites. Dramatic reductions in viscosity can occur in some highly filled liquid polymers Low friction particulate fillers when a relatively small quantity of an approParticulate fillers such as Teflon@powder, priate coupling agent is added to the molybdenum disulfide and graphite have formulation. been added to plastics in order to obtain low friction composites. Typical end uses have Silanes been oil-free bearings and molded plastics Silanes are currently the predominant couinserts for sliding drawers or windows. pling agents or surface treatments for particulate fillers. Many commercial fillers are available with various silane treatments and each silane is designed for optimum use in a specific type of polymer.R
Representative properties of particulate fillers Titanates and zirconates Titanates and zirconates produced by Kenrich Petrochemicals Inc., Bayonne, New Jersey, have received much attention in recent years and show promise of affording some remarkable improvements in processing characteristics and final properties of many composites systems. These coupling agents are analogous to the silanes. Each product has organic functionality and an inorganic backbone, so that one end can interact with the matrix resin and the inorganic component will have an affinity for the filler or reinforcement surface.
Miscellaneous coupling agents There have been many polymer additives and modified polymers that have been used to improve the properties of particulate filled polymer composite^.^ Polypropylene has been grafted with acrylic acid and this modified polypropylene has provided improved bonding to mica and talc in polypropylene composites. Copolymers of styrene and maleic anhydride have been shown to have improved mechanical properties in filled thermoplastics compared to a polystyrene matrix. Dupont produces a line of fluorosurfactants, with the trade name Zonyl, that have been recommended for use as a polymer additive to improve wetting and bonding with fillers. 11.5 REPRESENTATIVE PROPERTIES OF
PARTICULATE FILLERS 11.5.1 PROPERTIES OF PARTICULATE FILLERS
In considering the use of particulate fillers all factors should be evaluated including chemical and physical properties, surface characteristics, particle size and size distribution. Among physical properties are specific gravity, bulk density, specific heat, coefficient of linear expansion, index of refraction and Mohs' hardness value. These properties can
251
vary widely for the many different types and grades of particulate fillers. 11.5.2 OIL ABSORPTION NUMBER
The oil absorption number is an important characteristic of a particulate filler. An analogous situation exists in the paint and coatings industry, where two ASTM specificationshave been used to determine the oil absorption of pigments. The objectives of these tests are 'to obtain information about the vehicle demand of the pigment when it is used in a pigment paste.' In ASTM D1483, linseed oil is added in drops to the gently stirred pigment. The amount of oil required to form a paste is used to calculate an oil absorption value. In ASTM D281, linseed oil is added dropwise to a small quantity of pigment, while rubbing vigorously with a spatula. The end point occurs upon obtaining a 'stiff, putty-like paste, that does not break or separate.' ASTM D281 refinements, which can involve the use of appropriate liquids other than linseed oil, have been described for determination of oil absorption or packing fractions of fi1lers.l" In most applications, it is desirable to choose fillers with low oil absorption values, since these fillers will cause the minimum increase in viscosity of the matrix material. 11.5.3 HARDNESS
The particulate fillers that are commercially available have a wide hardness range. They are usually rated by use of qualitative scale in which the hardness of a mineral was determined by its ability to scratch or be scratched by another mineral. The scale was set up by a mineralogist, Friedrich Mohs. In the Mohs' ratings, the standard hardness minerals are: talc, 1; gypsum, 2; calcite, 3; fluorite, 4; apatite, 5; feldspar, 6; quartz, 7; topaz, 8; alumina, 9 and diamond, 10. The hardness can be important in the choice of a particulate filler. In the case of a polymer matrix cast part that must be post machined, it will usually be advantageous to
252 Particulatefillers choose a soft filler such as talc or calcium carbonate, since an alumina filler may result in severe abrasion of standard machining tools. Also, in the extrusion or injection molding of polymer composites, a quartz or alumina filler may cause severe abrasion of processing equipment or molds.
much higher ratio of resin/filler to fill the voids before the system becomes sufficiently fluid for most processing or molding procedures. Many applications will benefit from the use of a particulate filler and the proper packing of the filler, filler blends or filler and fiber combinations will be useful to optimize the physical properties of the composite.'l
11.5.4 TOXICITY CONCERNS
Extremely fine particle size fillers can present a health hazard and should be handled with special precautions. During recent years, fillers with a high content of fine crystalline silica have been withdrawn from the market as a result of concern as a suspected carcinogen. Other particulates with well known health hazards are asbestos and beryllium oxide powders. In general, most particulate fillers do not represent a significant hazard when handled by use of standard industrial safety procedures.
11.6.2 SHRINKAGE OF PARTICULATE FILLED PLASTICS
An important advantage provided by the use of particulate fillers is the reduction in the shrinkage of polymers, so that less warpage is obtained in large molded plastic parts. However, designers of molds for unfilled plastics must be careful to modify the design for use with particulate filled plastics. The lower shrinkage can result in parts that are out of tolerance, when mating structures are involved. Also, there may be problems in efficient removal of parts from a mold.
11.6 DESIGN CONSIDERATIONS IN THE USE
OF FILLERS 11.6.1 PACKING CONCEPTS
An understanding of packing concepts is important for optimum physical properties in most particulate filler composites. It is easy to visualize this concept in its simplest form, from the following considerations. When solid spheres of any uniform size are placed in a container, the spheres will occupy 62.5%) of the container volume and the interstitial void content will represent 27.5% of the volume. If much smaller spheres, with a size that permits them to fit within the interstitial space of the larger spheres, are now placed into the container, it is obvious that the interstitial void volume will be reduced. By appropriate choices of the sizes of the large and small spheres, a void volume of only 15% can easily be achieved with this bimodal packing. The difference between 27.5% voids and 15% voids can be extremely important in compounding with liquid polymers or resins, since the high void content requires a
11.6.3 ORIENTATION OF FLAKE FILLERS
The design of molds for flake filled polymer composites must include the flow orientation of the flakes. This can result in directional strength problems unless injection gate shapes and locations are properly designed. Another consideration for injection molding of these polymer composites is that flow conditions often result in the formation of knit lines, which may have much lower strength than the base material. 11.6.4 FILLER ADDITION REDUCES ELONGATION OF POLYMERS
The elongation at failure of a polymer will be greatly reduced when compounded with a particulate filler. This can lead to a significant reduction of toughness or impact resistance. This factor must be evaluated before making a final choice in the use of particulate fillers and loading levels.
References 11.7 TYPICAL END USES FOR PARTICULATE FILLERS
The traditional source for particulate filled plastics has been from plastics compounders. There are many custom compounders with appropriate equipment such as Banbury mixers and twin screw extruders that are used to obtain a uniform blend of the filler and matrix. Various filled polymer compounds, usually in the form of small pellets that have been chopped from extruded rods, are available from these compounders. In addition, large resin producers have introduced filled grades of their compounds. Among the many products that are commercially available are Nylon and PBT resins filled with talc or mica from GAF; Nylon filled with kaolin from HoechtCelanese, Dupont (Minlon), or Monsanto (Vydyne); and GE’s Valox 700, which is a PBT resin filled with mica. GE’s Valox HV7000 series resins are PBT composites that are highly loaded with mineral filler. Grade 7075 has 68 wt.% filler. Most commercial particulate filled plastics have had a maximum filler loading of about 40 wt.% because the physical properties have usually fallen dramatically when this level of filler has been exceeded. However, as the art and science of compounding progresses, it may be anticipated that future composites will effectively use much higher average loadings of mineral fillers. This trend is apparent from the high loading of minerals in GE’s HV7000 series mentioned above. Synthetic marble, which is used to fabricate bathroom sinks, is usually formulated by use of polyester resins that have been highly loaded with calcium
253
carbide fillers. Flexible and rigid plastic magnets contain about 90 wt.% loading of ferrite powder. These magnets are used in many automotive and appliance applications, such as door seals and refrigerator gaskets. Mica-reinforced polypropylene has been used in many automotive applications, such as fan shrouds, seat backs, glove compartment moldings and inner fender linings. Among the many end products that are made by use of particulate filled composites are electronic components, toys, marine components, grinding wheels, display items, tools, housewares and cameras. REFERENCES 1. Katz, Harry, S., and Milewski, John V., Handbook of Fillers for Plastics, New York: Van Nostrand Rheinhold (now Chapman & Hall), 1987. 2. Fillers and Extenders f o r Plastics, Norwalk, CT Business Communications Co., 1995. 3. Canova, L.A., Effect of surface treatments in mica-filled polypropylene, Plastics Compounds, 1990, July/August, 3843. 4. Davis, et d.,US Patent 2 460 977,1949. 5 . Searight, et d., US Patent 3 138 444, 1964. 6. Bland, et d., US Patent 3 150 947, 1964. 7. Schmidt, et d.,US Patent 3 190 737,1965. 8. Plueddeman, E.P., Silane Coupling Agents, New York Plenum Press, 1982. 9. Skeist, I., Handbook of Adhesives, New York: Van Nostrand Reinhold, 1989. 10. Ferrigno, T.H. in Handbook of Fillers for Plastics, New York: Van Nostrand Reinhold (now Chapman & Hall), 1987, pp. 17,19-20. 11. Milewski, J.V. and Katz, Harry S., Handbook of Reinforcements for Plastics, New York: Van Nostrand Reinhold (now Chapman & Hall), 1987, pp. 14-33.
SANDWICH CONSTRUCTION
12
Andrew C. Marshall
12.1 INTRODUCTION
This chapter covers a unique form of composites known as ’structural sandwich construction’. A structural sandwich consists of three elements, as shown in Fig. 12.1:
I
1. a pair of thin, strong facings; 2. a thick, lightweight core to separate the facings and carry loads from one facing to the other; and 3. an attachment which is capable of transmitting shear and axial loads to and from the core.
This chapter provides a general background and a brief summary of the various materials in common use; the design steps used to calculate loads; some design details for solving load point, edging and attachment problems; and a few tables, charts and graphs containing useful information for the designer. An attempt is also made throughout the chapter to provide suggestions and perspectives to help a new user of sandwich structures technology to avoid some of the errors of his predecessors. Structural sandwich construction is one of the first forms of composite structures to have attained broad acceptance and usage. Virtually all commercial airliners and helicopters and Fig. 12.1 The elements of a sandwich structure are nearly all military air and space vehicles make as follows: (a) two rigid, thin, high strength facings; extensive usage of sandwich construction. In (b) one thick, low density core; and (c) an adhesive recent years, most commercial space vehicles attachment which forces the core and facings to act as a continuous structure.The facings of a sandwich have also adopted this technology for many panel act similarly to the flanges of an I-beam, components. The effectiveness of sandwich resisting the bending loads and increasing the construction is shown in Fig. 12.2. bending stiffness of the structure by spreading the In addition to air and space vehicles, the sysfacings apart. However, unlike the I-beam’s web, tem is commonly used in the manufacture of the core gives continuous support to the flanges or cargo containers, relocatable shelters and airfacings. field surfacing, navy ship interiors, small boats and yachts, duplicate die models and production parts in the automobile and recreational Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Fncing material 255
Fig. 12.2 A striking example of how conversion to sandwich stiffens a structure without materially increasing its weight. This example uses 1.6 mm (0.063 in) thick aluminum facings and 1/4-5052 37 kg/m7 (2.3 lb/fPj aluminum core.
vehicle industry, snow skis, display cases, residential construction materials, interior partitions, doors, cabinets and a great many other everyday items. Although the employment of sandwich design to produce lightweight or special purpose load-carrying members is thought to have originated as early as 1820, routine commercial use of the idea did not occur until about 110 years later. What started this sudden acceptance was the successful commercial production of structural adhesives, starting in both UK and USA in the 1920s and 1930s. This early production began with the use of casein glue and later urea-formaldehyde and phenolics, with wood facings and cores. The search for better adhesives subsequently resulted in the development of the rubberphenolics and the vinyl-phenolics, which were suitable for use with metals. Commercial adhesives such as ’Cycleweld,’ (from Chrysler Motors), ’Plycosite,’ (from US Plywood) and ’Redux’ (from Bonded Structures, in Duxford, UK) adhered well to both wood and metals and possessed rather high and predictable strength. The result was the beginning of a revolution in bonding technology. Many further developments followed in only a few years. They included improved cleaning methods for metal skins; low weight, high strength/stiffness honeycomb core materials; ‘B’ staged tape adhesives which could be stored for long times; glass fabrics and collimated tapes preimpregnated with accurately measured
amounts of ’B’ staged resins; high strength resins; tough, high peel adhesives requiring lower cure temperatures and pressures; as well as the discovery of the resistance of sandwich to sonic fatigue. 12.2 FACING MATERIAL
The primary function of the face sheets in a sandwich structure is to provide the required bending and in-plane shear stiffness and to carry the edgewise and bending loads, as well as the in-plane shear loading. In the aerospace field, facings most commonly chosen are resin impregnated fiberglass cloth or a laminate of unidirectional fibers (commonly called ’prepreg’), graphite prepreg, 2024 or 7075 aluminum alloy, titanium alloy, or any of several stainless steel or refractory metal alloys. Even the most economical of these products represents a substantial cost and customary practice is to choose among them very carefully on a value engineering, or lowest lifetime cost, basis. 12.2.1 SUITABILITY OF MATERIALS
When choosing facing materials (as well as the core, adhesive, or other materials) for an application, it is wise to examine the less obvious properties of the material, such as toughness or brittleness, mode of fracture, durability and weatherability, compatibility with rivets and bolts and other such attributes which may directly affect the usability or success of the
256
Sandwich construction
end product, even though not directly involved in stress analysis or weight savings. An understanding of these requirements has resulted in a switch from aluminum to fiberglass skins and from fiberglass to aramid (Nomex, from DuPont) cores for most aircraft cabin interior panels.
applications such as boat hulls, large tanks and airborne pallets and containers. This broadening usage is also prompted by its excellent compressive strength and modulus properties when compared to all but the aramid paper honeycombs, which are much more expensive. Complete information can be obtained from the leading producer of these materials, BaltekI3,or Balsa Ecuador Lumber Company.
12.3 CORE MATERIALS
The primary function of a core in sandwich structures is that of stabilizing the facings and carrying most of the shear loads through the thickness. In order to perform this job efficiently, the core must be as rigid and as light as possible and must deliver uniformly predictable properties in the environment (such as high humidity) in which the finished part is to perform. 12.3.1 TYPES OF CORE MATERIALS
Wood Several different materials are used extensively as sandwich cores. The oldest of these is wood, which continues to be used in many applications as a core for such common applications as doors, partitions and many other ’builder’s supply’ items. It is also used in the majority of snow skis, either flat-grain or endgrain, although a few of the higher performance skis employ honeycomb, foam, or reinforced plastic cores. End-grain balsa has broad acceptance in boat hulls up to lengths of 15.2m (50 ft) or more and is still used for replacement flooring for many older and a few new aircraft. The traditional advantage of the low cost of wood has been progressively eroded with the passage of time and many users report difficulty in supply, even at prices higher than foam and sometimes approaching that of honeycomb. Even so, the ease of use and excellent durability of the end product has led to substantially increased usage, particularly of the carefully selected grades of end-grain balsa, in
Foam The use of foam as a structural core has been and is now, extensive. Recent developments in the technology of foam injection have sharply increased the use of these materials. The most novel of these is use of a cold-cavity die, in which the foam is injection molded in a single production step. A careful adjustment of the mixing and curing reaction of the foam, together with the heat-sink effect of the mold results in a part with facings which are simply an un-foamed, higher density form of the same polymer which constitutes the foamed core. The high productivity and modest cost of this scheme have resulted in many applications in the automotive and industrial fields. Another fast-growing form of the material is in cores for fiberglass snow skis and tennis rackets, in which an assembly of facings and close-out details is placed in a closed cavity mold and foam injected to form both the core and the adhesive attachment to the pre-cured glass fiber skins and various edge details. The saving in labor over conventional assembly methods has resulted in rapid acceptance of the process and the construction of many new factories. Foams can also provide special properties such as insulation or radar transparency, when used with appropriate facing materials. The very low cost polystyrene foams are used primarily in non-sandwich applications, their role in structural parts for refrigerated vehicles and buildings having been largely taken over by the urethanes. The single major
Cove materials 257 exception to this statement lies in the extensive use of polystyrene foams as cores in several thousand amateur-built composite aircraft. This application was pioneered by Burt Rutan, in his ’moldless construction’, used in his series of high performance small aircraft and the many similar designs offered by others in subsequent years. The polyvinyl chloride (PVC) foams, which made an impact on the transport aircraft industry as flooring cores, have been largely replaced by the more efficient high density aramid honeycombs. The foam-in-place system of producing sandwich structures has been used for more than 35 years, because of its simple concept. However, users of this system have always had difficulty with the continuing problem of producing uniform properties from one mix to the next and in achieving uniformly high core and bond strengths to the metal or pre-cured glass fiber skins. The use of systematic incoming inspection, automatic mixing and dispensing equipment and, in the case of critical airframe parts, test coupons, produced integrally with the basic part, have all helped to keep the problems under control. It will be noted that Table 12.1 does not list the shear strength of many of the various
foams, even though this value is needed for sandwich panel design. This property, even where listed, cannot be considered to be a reliable value. The actual value for an application at hand must be determined for the actual materials and conditions of use in order to be considered reliable. When a value for shear strength is not available, it may be roughly estimated to be about 0.7 times the compressive strength shown. Even the compressive strength cannot be considered to be reliable, however, as many differing methods of measuring this value are commonly used and each results in a substantially different value reported. 12.3.2 HONEYCOMB
Honeycomb types in common usage include products made from uncoated and resinimpregnated kraft paper, various aluminum alloys, aramid paper and glass or carbon fiber reinforced plastic in a number of cloth weaves and resin systems. Honeycombs based on titanium, stainless steel and many others are used in lesser quantities. Most honeycomb cores are constructed by adhesively bonding strips of thin material together, as shown in Fig. 12.3. In the case of aramid paper honeycomb, the
Roll
c
T
r
A
HOBE Block
HOBE Slice
+
Expanded Panel
Expansion Process of Honeycomb Manufacture Corrugated Sheet
Roll
Corrugating Rolls
Corrugation Process of Honeycomb Manufacture
I Corrugated Block
Fig. 12.3 Most honeycomb is produced by the expansion process. Actual cell shape produced by either method may vary greatly.
258 Sandwich construction Table 12.1 Properties of several foam materials used as cores* TYP Tensile strength (ASTM 01623)
Density
Compressive strength at 10% deflection (ASTM 01621)
lb/ft3
kg/m3
ABS (acrylonitrile bu tadiene-styrene) Injection molding type pellets 40-56
641-897
Cellulois acetate Boards and rods (rigid, closed cell foam)
6.0-8.0
96-128
170
1.2
125
Epoxies Rigid closed cell, precast blocks, slabs, sheet
5.0 10.0 20.0
80 160 320
51 180 650
0.35 1.2 4.5
90 260 1080
3-17 20-54 80-130
0.021-0.12 0.1384.372 0.552-0.896
2-15 22-85 158-300
Phenolics Foam-in-phase liquid resin
'X-1% 5-24 2-5 32-80 112-160 7-10
psi
MPa
psi
MPa
"F
"C
176-180
80-82
0.86
350
177
0.62 1.8 7.4
350 350 350
177 177 177
20004000 13.8-27.6 2300-3700 15.8-25.5
0.014-0.10 0.15-0.58 Continuous 1.09-2.07 service at 145 300
Polypropylene Pellets
50
801
5500
37.9
7500
51.7
Polypropylene"
35.0
561
1600
11.03
2100
14.4
Polyurethaneb
1.3-3.0 4-8 9-12 13-18 19-25
2148 64-128 144192 208-288
15-96 90-290 230450 475-700 775-1300
0.104.65 15-60 0.62-1.99 70-275 1.58-3.10 290-550 3.284.83 650-1100 5.34-8.96 1200-2000
100-2700
0.68-18.6 15-1500
1000 andup
6.90 andup
Skinned molded (rigid) Skin Core Polyvinyl chloride Rigid closed cell boards and billets
30p400
25-65 400-1041 3-30 48481
3
48
6
96
Maximum service temperature
270
132
0.10-0.41 0.48-1.90 1.99-3.79 4.48-7.58 8.27-13.8
180-250 200-250 250-275 250-300 250-300
82-121 93-131 121-135 121-149 121-149
40-3000 15-1500
0.28-20.7
150-250 66-121 150-250 66-121
95
0.65
200
1.38
* Where shear strength and modulus properties are not shown, use a figure of 0.7 times the compressive strength shown first approximation for design feasibility consideration. Always test actual material used for true value of shear strength and modulus. a High density, foam, molded, parts and shapes, with solid, integral skin. Rigid (closed cell) molded parts; boards, blocks, slabs; pipe covering; one-shot, two- and three-package systems for foam-in-place;for spray, pour, or froth-pour techniques. as a
Core materials 259 Table 12.1 Continued
Type
Thermal conductivity
Shear strength
BTU in h-'Pf2
Wm-' K-I
0.58-2.1
0.08-0.30
Cellulose acetate Boards and rods (rigid, closed cell foam)
0.31
0.04
Epoxies Rigid closed cell, precast blocks, slabs, sheet
0.26 0.28 0.32
0.04 0.04 0.05
ABS (acrylonitrile butadiene-styrene) Injection molding type pellets
Phenolics Foam-in-phase liquid resin
0.2 1-0.28 0.20-0.22 0.24-0.28
~ _ _
0.15
Polypropylene"
4.2
0.61
Polyurethaneb
0.11-0.21 0.15-0.29 0.19-0.35 0.26-0.40 0.34-0.52
0.2-0.4 0.02-0.04 0.03-0.05 0.04-0.06 0.05-0.07
0.12-0.80 0.21-0.55
0.02-0.12
2.0 at 70
MPa
psi
MPa
20 90 180
0.14 0.62 1.24
226 1500 4500
1.56 10.3 31.0
450
3.1
15000
103.5
0.03-0.04 0.03-0.04
1.05
Polyvinyl chloride Rigid closed cell boards and billets
psi
0.03-0.04
Polypropylene Pellets
Skinned molded (rigid) Skin Core
Shear modulus
225-15 000
20-500 65 120
0.45 0.83
1200 2200
8.3 15.2
* Where shear strength and modulus properties are not shown, use a figure of 0.7 times the compressive strength shown as a first approximation for design feasibility consideration. Always test actual material used for true value of shear strength and modulus. a High density, foam, molded, parts and shapes, with solid, integral skin. Rigid (closed cell) molded parts; boards, blocks, slabs; pipe covering; one-shot, two- and three-package systems for foam-in-place; for spray, pour, or froth-pour techniques.
260
Sandwich construction
inherent toughness and abuse resistance of the enced by the properties of the materials from material makes cores of 1 6 4 8 k g / m 3 (1-3 which they are manufactured. Some of these lb/ft3) an excellent choice for aircraft cabin differences are obvious in the thermal conducinterior walls and ceilings, even with glass fab- tivity information shown in Fig. 12.4 and Fig. ric-reinforced skins as low as 0.254 mm (0.010 12.5. However, several significant properties of honeycomb cores are peculiar to the materiin) in thickness. Physical and mechanical properties of the als and should be separately noted. honeycomb core materials are strongly influThermal Resistance - Aluminum Honeycomb
.028 N
E13021 W
P-=x
&j.014 w 0) U
007
(4
0 25(1 0)
5 0 (20)
lO(40)
76(30
Core Thickness- cm (in )
Thermal Resistance-Non Metallic Honevcomb
Fig. 12.4 Thermal conductivity through sandwich panels can be isolated into the contribution of each component: facings, core and adhesive. The resistances (or reciprocal of conductivity) can simply be added - including the effect of boundary layer conditions. The thermal properties of typical facing materials may be found in many handbooks. Thermal resistance values for typical core to facing adhesives are typically 0.03 for film adhesives with a scrim cloth support and 0.01 for unsupported adhesives. These graphs give the resistance for aluminum and non-metallic honeycomb at a mean temperature of 23.9"C (75°F). Note that for non-metallic honeycomb, it has been found that the cell size is more critical than core density. The reverse is true with aluminum honeycomb. To correct for mean temperature, divide the resistance at 23.9"C (75°F) by coefficient Q.
70
4
53
3
35
2
18
1
cu
$13 W
9 2 e
PI U
1 3 (05)
5 0 (20 )
3 8(1 5)
25(10)
Core Thickness- cm (in )
Effect of Mean TemDerature
-1 29
-17.8
93
204
Core materials 1.2
,
I
I
I
I
I
I
1
I
I
I
I
I
I
261
I
1.1 Y
a
8
1.o
.9
.e
Y 4:
5 F 8 0 0
.7 .6 .5
.4
I
I
Fig. 12.5 Measured core shear strength will vary depending upon the test method, core thickness, skin thickness and many other factors. The above curves may be used only for preliminary correction factors. Physical tests of the final design must be used to confirm actual values obtained, as the curves shown above are only approximate.
1 (bl 0
I'
125 L E KRAFT PAPER, PHENOLIC RESIN, & " I N C H XEXAGON CELLS 3003 - H I 9 ALUMINUM 00024 -INCH FOIL ;':INCH HEXAGON' C E L L S
I I I I 2 3 CORE THICKNESS (INCHES J
1 I 4
Density All mechanical properties increase with higher density, as shown in Fig. 12.6.
Cell shape All honeycombs are anisotropic and the resulting directional properties should be adapted to
the loads anticipated. Figure 12.7shows typical differences in shear strength for the L and W directions. In addition, some cell shapes allow easy forming or curving at a small loss in strength/weight ratio. This attribute can be of great importance in manufacturing curved parts of appreciable thickness.
Fig. 12.6(a) Typical stabilized compressive strengths.
262 Sandwich construction
1 PCF I
0
20
40
60
80
100
120
140
L
160kglm'
Density
Fig. 12.6(b) Typical 'L' shear strengths.
8
Fig. 12.7 Plate shear test values may be significantly different from test results obtained from testing beams. Values shown above are typical for 5052 aluminum honeycomb.
Cell shape variations Cell shape variations may be either furnished to specification by the core manufacturer, or, in certain materials such as aluminum, shapes may be intentionally or inadvertently altered by the core user. It should be noted that the under- or over-expansion of the core changes its cell shape and density. The over-expanded version of Fig. 12.8(c) changes directional properties such that the L direction becomes slightly the weaker of the two major axes.
C
D
E
F
G
H
Fig. 12.8 A few of the many cell configurations in common usage. 8, G and H are only produced by the corrugating method. F is a cell configuration nearly always used in the manufacture of welded metal honeycomb. C is flexible in one axis, while G and H are flexible in both axes. A, C and D are expanded from identical unexpanded slices, A being normal expansion, C fully over-expanded and D 50% expanded. B is a reinforced corrugated core, with an extra layer of uncorrugated web material placed between each layer of corrugated web material. Reinforcing layers may be added in discrete locations or patterns and may be of the same or different web material or thickness.
Core materials
263
Since the drop in the L direction strength can amount to as much as 30%, such changes in cell shape must not be allowed to occur by error.
Cell size Although cell size tends to be a secondary variable for most mechanical properties of core materials, it is primary in fixing the strength level of the core-to-face attachment (or, more accurately, in fixing the required lower limit on core-to-panel adhesive weight) and in determining stress levels at which intracell buckling or face dimpling of facings occurs.
Thickness The shear and compressive properties noted for a 'pecific 'Ore type can Only be 'Pecified and when test methods are controlled and the correct thickness of core is tested. Failure to allow for the effect of thickness can affect observed values by a factor of 4 or more, as noted in Fig. 12.5. It should be emphasized that the correction factor shown may be considerably different, depending on skin material and thickness, as well as the exact test method used.
v :
Fig. 12.9 Plate shear test for honeycomb shear strength and modulus 1.27 cm (0.50 in) thick steel plates are oven-cleaned and may be reused many times.
Specimen geometry and test method Like thickness, these must be specified and carefully controlled in order to obtain comparability -with test values obtained by others. Shear strength values obtained using plate shear test methods of Fig. 12.9 are quite normally up to 25% below those obtained when using the flexure method shown in Fig. 12.10. Both methods are accepted and used and any lack of understanding of the differences can lead to monumental, if nonsensical, problems. It will be noted that the tables of mechanical properties for various honeycombs, Tables 12.4-12.12, specify the shear test method used in producing the data shown.
Fig. l2*l0 Short beam shear test for core. Note the ample bearing area provided at each load and support point to preclude core crushing ,-hear failure. prior to
Paper honeycomb
Paper honeycomb is the first predecessor of all the types of honeycomb, having been produced for some 2000 years. Early forms were not used as structural cores, but were
264 Sandwich constructiotz
Core materials 265
266
Sandwich construction
. -
m n
x x x x 0000 \ \ \ \
yE z2 z2 z2 z2 6
a,
L
3
Core materials 267
268 Sandwich construction employed as decoration - and are still frequently seen today as seasonal decorations in department stores in the form of expanded bells, spheres and so forth. Current materials used as sandwich cores are different, in that much stronger kraft paper is employed and 11-35% phenolic resin is frequently used to improve mechanical properties, as well as moisture and fungus resistance. Many variations are available in cell sizes of 10,13 and 19 mm (%, X and % in) or even larger sizes. The higher strength versions are only produced in the smaller cell size, with the 10mm (% in) cell available as a watermigration resistant grade meeting military specification MIL-H-2104Q. Most applications are found in non-aircraft uses, where cost saving is the one primary objective. Usage is growing rapidly in recreational vehicles; for doors, walls and partitions; for factory produced kitchen cabinets; in packaged patio room additions for homes; in curtain wall panels; and in bearing walls for commercial building.
Some of the above alloys are also available as corrugated, corrugated and reinforced, overexpanded and flexible cell configurations. Some have also been produced in a specially tailored geometry to make all the cell axes lie on a true radius of a cylinder, a sphere, or other unique configurations. These same alloy foils can also be wound as a corrugated spiral to form a cylinder or tube for very light energy absorption applications. The aluminum honeycomb cores remain the most used, as well as the most versatile of the various core materials obtainable and are often found to possess the most favorable performance/cost ratio available. Expanded aluminum cores commercially available ranges from a low of about 32 kg/m3 (2 lb/ft3) to a high of 192kg/m3 (12.0Ib/ft"). Corrugated aluminum cores, however, start at under 128kg/m3 (81b/ft3) and can be purchased up to 880 kg/m3 (55 lb/ft3). At densities below 128 kg/m3 (8 lb/ft3) corrugated core suffers a serious penalty in shear properties when compared to expanded core.
Aluminum honeycomb
Glass fiber-reinforced plastic honeycomb
This family of materials has been in production and growing since about 1947.Aluminum honeycomb now includes four alloys, at least five cell shapes and many foil gauges to provide a range of densities. The alloys generally available are:
This family of materials is most commonly used in electrically sensitive parts, such as radomes and antennae, or where a heat resistant resin and low thermal conductivity make it a natural choice. It has also seen distinguished service as a matrix for retaining non-structural ablative materials, such as soft silicone rubbers or syntactic rigid epoxy foams, which otherwise could not have been used effectively as ablative heat shields on the Gemini and Apollo re-entry vehicles. Only high temperature phenolic and polyimide cores are generally produced. They are commonly available in cell sizes of 5, 6.3 and 10 mm (K, X and X in) with a 3 mm (% in) cell available in a bias weave glass reinforcement. Densities range from 32 to 192 kg/m" (2 to 12 lb/ft3). Mechanical properties of several commercially available glass fiber-reinforced cores are shown in Tables 12.3-12.6.
0
0
0
0
3003-H19, the lowest strength of the group, usually used for non-aircraft applications; 5052-H39, the most often used aircraft grade, available with a corrosion resistant surface treatment. Mechanical properties are listed in Table 12.2; 5056-H39, the strongest of the regular aircraft grades, available with a corrosion resistant surface treatment; 2024-T3 or T81, the most heat-resistant alloy and slightly stronger in some properties than 5056-H39. Available with a corrosion resistant surface treatment.
Core materials 269 Table 12.4(a) Properties of glass-reinforced phenolic honeycomb (bias weave reinforcement)*
Conipressiue
Honeycomb designation material - cell - density__ HFT - 1/8 - 3.0 HFT - 1 / 8 - 4.0 HFT - 1 / 8 - 5.5 HFT - 1/8 - 8.0 HFT - 3/16 - 1.8 HFT - 3/16 - 2.0 HFT - 3/16 - 3.0 HFT - 3/16 - 4.0 HFT/OX - 3/16 - 6.0
Bare Strength, psi
Plate shear
~~-
-~_____.
Stabilized Strength, Modulus, psi ksi
-.--_____ ~
-.
'L' Direction
~
'W' Direction
__
Strength, psi
Modulus, ksi
Strength, Modulus, psi ksi
typical
typical
typical
typical
typical
typical
typical
300p 390p 52533 1450p 75P
350p 575p 960p 1625p 120p 170p 375p 550p 11OOp
22p 45p 67p
185p 300p 425p 575p 105p 115p 200p 275p 290p
17P 32P 42P
95P 150p 225p 340p 5% 6OP
7P 12p 17p 25p
loop
9P
140p 335p
30p
loop 27513 435p lOOOp
1OOp 14p 17p 32p 45p 67p
48P 13P 15P 24P 3% 13P
4P 5P 12p
* Test data obtained at 0.500 in thickness. Honeycomb is normally not tested for bare compressive strength.
Table 12.4(b) Properties of glass-reinforced phenolic honeycomb (bias weave reinforcement)*(metric) ~ _~_
Honeycomb designation material - cell - density HFT - 1 / 8 - 3.0 HFT-1/8-4.0 HFT - 1 / 8 - 5.5 HFT - 1 / 8 - 8.0 HFT - 3/16 - 1.8 HFT - 3/16 - 2.0 HFT - 3/16 - 3.0 HFT - 3/16 - 4.0 HFT/OX - 3/16 - 6.0
Bare Strength, kPa
Compressive
Plate shear _____
Stabilized Strength, Modulus, kPa MPa
'L' Direction
Strength, kPa
Modulus, MPa
'W' Direction Strength, Modulus, kPa MPa
typical
typical
typical
typical
typical
typical
typical
2068p 2688p 3619p 9997p 517p 689p 1896p 2999p 6894p
2413p 396413 6618p 11203p 827p 1172p 2585p 3792p 7584p
151p 310p 461p 689p 97P 117p 220p 310p 461p
1275p 206813 2930p 3964p 724p 792p 1378p 1896p 1999p
117p 220p 289p 331p
655p 1034p 1551p 2344 344p 413p 68913 965p 2309p
48P 82P 117p 172p 27P 34P 6% 82P 206p
89P
103p 165p 206p 89P * Test data obtained at 12.70 mm thickness. Honeycomb is normally not tested for bare compressive strength.
Aramid paper honeycomb This is an especially tough and damage resistant product, based on a completely synthetic, calendered 'Nomex' paper material produced by DuPont. The core is expanded very much like aluminum or glass fabric honeycomb and then dip-coated with phenolic or other suitable resin system. The mechanical properties of the material as a structural core are some-
what lower than aluminum, especially in modulus, but it possesses a unique ability to survive overloads in local areas without permanent damage. This translates into abuse resistance when applied to very light interior aircraft panels or flooring and gives the material a competitive edge even at the higher cost it represents. The base material is relatively incombustible and the small amounts present
270 Sandwich construction Table 12.5 HFT glass-reinforced phenolic honeycomb (Fibertruss bias weave)* ~~
_
Honeycomb drsignation, materid - cell - densitu
HFT - 1/8 - 3.0 HFT - 1/8 - 4.0 HFT - 1/8- 5.5 HFT - 1/8 - 8.0 HFT - 3/16 - 2.0 HFT - 3/16 - 3.0 HFT - 3/16 - 4.0 IIFT/OX - 3/16 - 6.0
~
Compressive
_
Bare Strength, mi
_
~
_
_
_
_
Stabilized ~ _
.
_
Strength, Modulus, asi ksi
_
~~~~
~~
_
~
~
_
Plate shear
_
_
'L' _Direction
Strength, psi
Modulus, ks i
~
__-
'W' Direction-~ Strength, Modulus, psi ksi
typical
typical
typical
typical
typical
typical
250p 460p 850p 1600p 90P 250p 460p l000p
360y 530p 9501) 1750p 140p 320p 530p 1100p
21P 45P 65P 95P 17P 32P 45P 67P
185p 310p 460p 600p 118p 170p 310p 290p
16P 25P 34P 43P 15p 20P 25P 13P
96P 150p 240p 340p 55P 90P 150p 335p
typical 6.4~ 9.5p 13.5p 20.0p 4.3p 6.5~ 9.5p 30.0~
* Test data obtained at 0.500 in thickness. p = preliminary properties
in typical panels result in low volumes of smoke and gases given off in fire tests. Typical applications make use of these properties very effectively.As a consequence, they have grown to a commercial volume nearly as large as that of aluminum, for use in aircraft structures. Uses outside the aerospace industry are limited due to the high cost of the material, but despite this it has seen some application in boat hulls up to 10.2 m (40 ft) in length, as well as in skis, racing shells and several other products. Aramid core is normally produced in cell sizes of 3, 5, 6.5 and 10 mm (%, 36,X and % in), in densities of 24-192 kg/m3 (1.5-12 lb/ft"). Densities higher than 64 kg/m3 (4 lb/ft3) are almost entirely used for aircraft flooring. Mechanical properties of some of these core materials are shown in Table 12.6.
Carbon fiber honeycomb Reinforced plastic honeycomb has for many years employed glass fabric reinforcement,b u t only rarely employed other fibers. In the past few years, however, both Kevlar and carbon fiber have become much more common as reinforcing fibers for honeycomb. Carbon fibers only now are beginning to be used in
space vehicles. In addition to this small usage, however, carbon fiber honeycomb is now used as the structural core for nacelle assemblies in the Boeing Model 777 transport aircraft. The constant pressure for lighter structures in such designs has led to the use of carbon fiber facings, which have a potential corrosion problem when used with aluminum cores. This concern for corrosion problems has subsequently led to the adoption of a new class of carbon fiber honeycomb materials for this aircraft and will possibly lead to further use in other future designs. Two types of carbon fiber cores are now being produced. One is for purely structural applications, while the other has a requirement for heat transfer through the thickness of the panel. The former type uses only the usual pan based carbon fibers, while the latter employs pitch based carbon fibers, which duplicate the heat transfer properties of the aluminum core which it replaces. Although neither of these materials is as yet in large volume production, the economic impact is substantial, since these honeycombs are markedly higher in price than the aluminum or Nomex cores they replace. Little data is yet available on these new cores, but it is likely they will see substantial
Adhesive materials 271 use and public scrutiny in the next several years.
Kevlar honeycomb This honeycomb has been in use for a number of years as a core for space vehicle antenna reflectors. The purpose of the Kevlar honeycomb is to allow transmission of radio signals through the panel, while at the same time the Kevlar facing acts as a partial reflecting antenna for a different wavelength of a different signal. Kevlar honeycomb, based on one of several fabrics woven from Kevlar yarn, is usually produced in cell sizes of 6.3-9.5 mm (%-% in) . Usual densities available range from 16 to 64 kg/m3 (14 Ib/ft3).
be understood by the designer and fabricator in order for the otherwise inevitable problems to be avoided. Some factors which merit attention are discussed below. 12.4.1 PRODUCTS GIVEN OFF DURING CURE
Some adhesive types, such as phenolic, give off a vapor as a product of the curing reaction and the presence of these secondary materials can lead to several problems: 0 0
0
0
Kevlar paper honeycomb In addition to Kevlar honeycomb made from woven fabric, DuPont has recently introduced a new honeycomb, based on a Nomex-like paper, which is entirely composed of fibers derived from Kevlar. This material has rather surprising mechanical and physical properties, with strengths well above both glass and Nomex honeycombs and dielectric properties somewhat superior to Nomex. This material is trade named 'Kortex' and is available in the usual range of cell sizes and densities. Because the material is somewhat more expensive than Nomex, no large scale replacement of Nomex honeycomb appears likely, although many special purpose applications have been developed in both air and space craft. 12.4 ADHESIVE MATERIALS
Adhesives, as they apply to sandwich structures, constitute a somewhat different family of materials than those required to bond an open cellular core to a stiff and continuous facing. Although these differences are less important with some of the newer modified epoxy materials, they remain basic and must
internal pressure, resulting in little or no bond in some areas, or 'blisters'; core splitting, as the gas forces its way through the core to a lower pressure area; core movement, sometimes several inches, resulting in an unusable cured part; subsequent corrosion of core or skins by the chemical action of the vapor or its residual condensate.
12.4.2 BONDING PRESSURE
Adhesives such as the phenolics and some others actually require more than atmospheric pressure in order to prevent excessive porosity. Certain forms may be suitable for solid cores like balsa, but cannot be used at all in open cores such as honeycomb or large cell foams. Also, most core materials will not alone withstand compressive bonding loads exceeding a few atmospheres and consequently cannot be used with any adhesive system requiring higher pressures. 12.4.3 FILLET FORMING
In order to achieve a good attachment to an open cell core, such as honeycomb, the adhesive must have a unique combination of surface tension, surface wetting and controlled flow during early stages of cure. Controlled flow prevents the adhesive from flowing down the cell wall and leaving a low strength top skin attachment and an overweight bottom skin attachment.
272 Sandwich construction
Loloo
m m m
OOLo
+ e m
m a m
s; .N
6
2 -..
u u m
m o o d N N
o m 0 d m N m a -
Adhesive matevials 273
274 Sandwich construction 12.4.4 ADAPTABILITY
The requirements noted above must all be met while also meeting all the requirements of a skin-to-skin to skin-to-doubler attachment. In the case of contoured parts, the adhesive must also be a good 'gap-filler ' without appreciable strength penalty, since tolerance control of details is much more difficult to achieve on contoured than on flat panels and a greater degree of latitude for misfit must usually be allowed. 12.4.5 BOND LINE CONTROL
This is a need which exists because of misfitting details and is approximately the opposite of adaptability. It is the capability of the adhesive to resist being squeezed out from between faying surfaces when excessive pressure is applied to a local area of the part during cure. Many adhesives are formulated to achieve good core filleting and are subsequently given controlled flow by adding an open weave cloth or fibrous web, cast within a thicker film of adhesive. This 'scrim cloth' then prevents the faying surfaces from squeezing out all the adhesive, which would result in an area of low bond strength. 12.4.6 TOUGHNESS
The word 'toughness' has many meanings in the world of adhesives. Usually, it refers to the resistance shown by the adhesive to permitting bond line cracks to grow under impact loading. In the area of sandwich core-to-facing bonds, it refers to the resistance shown by the adhesive toward loads which act to separate the facings from the core under either static or dynamic conditions. It has been found from experience that greater toughness in the bond line usually equates to greater durability and thus to longer service life. Many types of tests have been devised to measure toughness, but the most common one used for sandwich structures is the climbing drum peel test (Fig. 12.11). This test has the
Fig. 12.11 Climbing drum peel test for adequacy of skin adhesion. The difference in diameter of the cylinders to which the straps are attached and the cylinder to which the skin is attached causes the drum to rotate clockwise when tension is applied by the universal testing machine. This arrangement allows duplication of test results from one shop to another.
virtue of being easily duplicated, as well as possessing an obvious relationship to the toughness whose value is sought. Values of peel strength will vary considerably, dependinn upon:
0 0
toughness of the adhesive; amount of adhesive used; density of the core; cell size of the core; direction of the peel (with or across the ribbon direction); adequacy of the surface preparation; degradation of the adherend surface subsequent to bonding.
Because these variables can lead to widely differing peel strengths for the very same adhesive, all of them must be properly understood and controlled if the peel test is to be used and its value compared to other test results.
Adhesive materials 275 The peel test is used to control quality throughout the sandwich industry. Values obtained, provided the adhesive weight and core material are in balance, will give indications of tooling or cure problems and of adherend surface preparation problems. It is particularly useful for this when an environmental exposure involving both elevated temperature and high humidity is interposed between manufacture and test. It is also adaptable to use with nearly any skin material, except that it becomes impractical with very thick or very stiff skins. It can be readily seen that a number of points of difference separate the sandwich adhesives from other structural adhesives. Fortunately for the sandwich user, many adhesives are available which satisfactorily meet both sets of requirements. me types available, along with some salient features, are as follows. 12.4.7 PHENOLICS BLENDED WITH VINYLS, RUBBERS OR EPOXY
All of these families of adhesives give off at least some water during cure and are therefore used only where their high strength, durability or high temperature mechanical properties are essential. Since the out-gassing cure products usually require venting or perforating the core material and a number of non-outgassing, high temperature adhesives have become available, their use as sandwich adhesives has sharply declined in recent years. 12.4.8 EPOXIES MODIFIED WITH NYLON OR OTHER POLYAMIDE POLYMERS
These adhesives were the first to have excellent filleting and controlled flow along with both high strength and high toughness, although they are somewhat moisture sensitive. Some versions are provided as one side of a two-sided tape adhesive, in which the other side is a rubber or vinyl-phenolic, to provide both excellent peel and durability at the skin side with excellent peel at the core side.
12.4.9 NITRILE RUBBER MODIFIED EPOXIES
These make up a broad group of more recent materials which provide much of the flow and toughness shown by the nylon-epoxies, along with the durability and weather resistance of the vinyl-phenolics. They are the most common of the 'toughened' thermosetting adhesives and are usually limited to about 149°C (300°F) service temperature. Some of these materials routinely achieve shear strengths of 34500 kPa (5000psi) and most can be cured over a wide range of temperatures and pressures. 12.4.10 URETHANES
Urethane based adhesives are used in commercial structures. Both moisture-cured and two-part systems are available. 12.4.11 OTHER POLYIMIDES, THERMOPLASTICS AND HIGHLY SPECIALIZED ADHESIVES
These are used in a number of applications ranging up to about 371°C (700°F)service temperature, but do not represent either a very large group of materials or a large volume of usage. In addition to categorizing the available adhesives by chemical type, they can be grouped by the form in which they are available. Generally these are as follows.
Light liquids, heavy liquids, thixotropic liquids, pastes, putties, or syntactic foams Only a few are used as a core-to-facebond, but many such materials are used in sandwich construction to splice pieces of core to each other in order to provide high strength edges, areas, or surfaces, or to carry shear loads from fittings, inserts, or end ribs. Most of the materials so used are epoxies, modified epoxies, epoxy polyamides or epoxy polyimides. Curing temperatures vary from as low as 4.4"C (40°F) for some two-part systems up to
276 Sandwich construction 216°C (420°F) for some of the materials intended for service at elevated temperatures.
All the above forms of adhesive are in current use at substantial volume and most are available from many sources.
Supported films Films or tapes having a carrier of light glass fiber, cotton, nylon, or polyester fabric, or spunbonded synthetic fiber are provided either dry or with slight to moderate ’tack’ or stickiness, so that the parts of the assembly stay in place as they are being assembled. Unsupported films, containing only the adhesive The very low weight films are nearly always furnished without a carrier, as the weight of the carrier itself becomes quite appreciable in very light sandwich structures. They are often hard to handle and sometimes have bond line control problems. Reticulating films These are intended for use at very low weights, with the adhesive being melted by hot air after placing on the core, so that it draws back to the cell edge and provides material to form the largest possible fillet without wasting any on the inside facing surface in the middle of the cell. Cell-edge adhesive This is a material pre-placed on the cell edge by the honeycomb manufacturer to provide the same results as those produced with reticulating films. Self-adhesive skins These skins are usually structural fabrics of glass, graphite, quartz, or aluminum coated glass fibers, pre-impregnated with a resin, which is then cured so that the fiber-filled resin becomes both the face structure and the attaching material.
12.5 DESIGNING A SANDWICH
The usual objective of a sandwich design is to save weight or to increase stiffness or to use less of an expensive skin material, or perhaps all three. Sometimes other objectives, such as reducing tooling or manufacturing costs, achieving aerodynamic smoothness, reducing reflected noise, or increasing durability under exposure to acoustic energy, are also involved. The designer’s problems sift down to relatively few, such as getting the loads in, getting the loads out and attaching small or large load-carrying members, under constraints of deflection, contour, weight and cost.
Understand the fabrication sequence and methods. The cost of a sandwich structure is fundamentally fixed at the design stage and a considerable difference in cost can result from alternate solutions to the design problem. Both of the edge close-out details shown in Fig. 12.12 perform essentially the same job at the same weight. Placing the legs of the channel facing outward instead of inward saves the cost of two relief cuts into the core and the very difficult step of sliding the edge of the core and adhesive into the channel. Another alternative at even lower cost for either fixed or simply supported edges is shown in Figs. 12.13-12.16. Use the right core. Several densities of core can be used in a single panel, each appropriate to the load carried in the area and adhesively bonded to its neighbor, as shown in Fig. 12.17. In many cases, however, the weight saved in lower density areas of core is added back in the form of core splice adhesive weight. Core splices, such as those shown in Fig. 12.18(b)or (c), have been used to produce ablative matrix structures for large re-entry heat shields,
Designing a sandwich 277
1
(
GOOD ~
..
POOR
.
Fig. 12.12 The square edge close-out shown here using a channel may result in a neat, clean edge, but requires machining both the top and bottom of the core and squeezing adhesive and core into the channel during assembly. The alternative shown on the left would be much better.
/
Densified Core 2
Fig. 12.13 Densified core edge treatment.
DENSIFIED CORE
HIGH -S TR€NG TH INSERT'
ME TAL CHANNEL
'
EDGE CELLS FILLED'
'
FACINGS FORMED
'
HIGH-STRENGTH INSERT,
ME TAL CHANNEL
'
HIGH - S TRENGTH INSER T\
METAL ''2
Fig. 12.14 Several common edge treatments.
CRUSHED AND BONDED r RE/NFORCEMENT
F O RMEO RING
---.
278
Sandwich construction Strong. Special extrusion. Seals can be incorporated.
Strong. Uses standard angle. Pop rivets to locate and apply pressure during bonding.
Very strong. Special extrusion. Difficult to apply adhesive uniformly and assemble Very strong with inside tie-bar. Can include external seal or gasket.
Low strength. Very low cost. Inside facing and core scarfed then bent. Fill corner with epoxy or foam to stiffen.
Fig. 12.15 Several suggestions for corner designs, edge close-outs and splices.
f
Ex trurion, Weldd
b \
Locking Bar
Fig. 12.16 Additional joints and corner treatments.
Fig. 12.17 Typical core splice using a foam-tape adhesive. Foaming of the tape adhesive permits a less-than-perfect fit of core details, but requires that the core be precisely fixed in position during the cure to avoid a step in the surface at the splice line.
Designing a sandwich 279 ACROSS RIBBON 1Wl DIRECTION
IN RIBBON /LI DIMCTION
C
Fig. 12.18 Joint A may be formed by simply crushing one piece of glass fabric honeycomb into the adjoining section. This method will work to some extent with some aluminum honeycombs, but not with most other core materials. Joints B and C require a perfect match of cell shape and cell pitch and are very difficult to produce on any realistic and cost-effective basis.
but become prohibitively expensive to produce for splices more than a few inches long. Do not hesitate to use several joining methods in the same part. Fittings to be included in a bonded sandwich may be produced from weldments, forgings or riveted assemblies, or may themselves be bonded assemblies. Available adhesives permit secondary bonding to be performed at temperatures from 16°C (60°F) up to 177°C (350°F) without degrading the integrity of the previously bonded sub-assemblies. Use bolts and rivets for carrying loads (not for soothing fears). Where space is not available for progressive doublers or wide-area bonded overlaps to carry high loads, the addition of rivets or bolts is sometimes the only solution. Their use, however, often results in lower (sometimes dramatically lower) fatigue life of the structure, in addition t o increased weight. The use of 'chicken rivets', added for the sole purpose of appearance, is to be particularly avoided, since they often defeat much of the advantage which would otherwise result from use of the bonded structure.
5. Use doublers where needed, instead of a heavier facing over the entire part. The use of doublers, although adding labor cost in assembly, often improves the part quality. Where skins are formed of glass or graphite prepreg, the problem is even simpler, since extra plies can be added to carry extra loads exactly where and as needed. 6. Use external doublers rather than infernal doublers wherever possible. The use of internal doublers usually means that a relief cut must be made in the thickness dimension of the core to prevent bridging and a consequent unbonded area where the doubler ends. Figure 12.19 shows a panel where the loads which can be carried are the same at each end of the panel. The design detail on the left end can cost substantially more to manufacture than that on the right end. Figure 12.20 shows the same panel with both ends produced at low cost, while still achieving an unbroken outer skin line on one side. In the case of some skin materials, such as 0.25mm (0.010in) aluminum, or most weights of pre-impregnated glass or graphite cloth, it is feasible to use thin doublers without a relief cut in the core, since
280 Sandwich construction
Fig. 12.19 Internal and external doubler treatment.
I
1
Fig. 12.20 Low cost doubler treatment.
Fig. 12.21 Doublers at a skin splice.
the gap caused by bridging is small enough to be within the capacity of the adhesive to fill. Sometimes an extra layer of adhesive film is added to help. An example of a double skin splice using this method is shown in Fig. 12.21.
12.6 STRUCTURAL ANALYSIS FOR SPECIFIC CASES
The following notations are used in sandwich design formulas.* This chapter’s formulas are only for honeycomb beams and columns which have the same facings on each side of the core.
Structural analysis for specific cases 281 D = flexural stiffness; D =
E t h2
JL-
2% Ec = modulus of elasticity of the honeycomb, Pa (psi) E, = modulus of elasticity of facing material, Pa (psi) Gc = shear modulus of rigidity of the honeycomb, Pa (psi) Kb = bending deflection constant Ks = shear deflection constant L = beam span length or column height, mm (in) b = beam width, mm (in) M = maximum moment, kg/m width (Ib/ in width) P = load, kg (lb) Pcr= column critical load, kg/m (lb/in) P = column facing yield load, kg/m (lb/in) = maximum shear force, kg/m width (Ib/ in width) d = sandwich total thickness, mm (in) h = distance between facing centroids, mm (in); h = tc + t, s = core cell size, mm (in) tc = core thickness, mm (in) t, = facing thickness, mm (in) w = uniform beam load, Pa (psi) A = maximum beam deflection, mm (in) ,I6 = 1- Poisson’s ratio of the facing material squared = 1- p 2 p = facing material’s Poisson’s ratio of = maximum facing stress, Pa (psi) uy= yield stress of facing material, Pa (psi) zc< = maximum core compressive stress, Pa (psi) zcs= maximum core shear stress, Pa (psi)
9
12.6.1 DESIGN REQUIREMENTS
Sandwich structures should be designed to meet the basic structural criteria listed below, when these criteria pertain to the type of loading under consideration. 1. The facings should be thick enough to withstand the tensile, compressive and shear stresses induced by the design load. 2. The core should have sufficient strength to withstand the shear stresses induced by the design loads. 3. The core should be thick enough and have sufficient shear modulus to prevent overall buckling of the sandwich under load. 4. Compressive modulus of the core and the compressive strength of the facings should be sufficient to prevent wrinkling of the faces under the design load. 5. The core cells should be small enough to prevent intracell dimpling of the facings under design load. 6. The core should have sufficient compressive strength to resist crushing by design loads acting normal to the panel facings or by compressive stresses induced through flexure. 12.6.2 MODES OF FAILURE
Typical modes of failure are shown in Fig. 12.22. 12.6.3 DESIGN STEPS
1. Define Zoads. For multi-point loadings, use the formulas in Roark’s Formulas for Stress
and Strain.16 The need to be able to accurately calculate the exact performance for many forms of sandwich structures had led to the development of a substantial body of literature on the subject. This chapter will cover in detail only the commonly used analyses and will provide reference sources for a number of others.
2. Define beam type. The values of Fig. 12.23 provide the simple starting point for these calculations. Some care in using the fixed end type of support is needed, as in actual practice total fixity is not realized and the resulting deflection is greater than that calculated. 3. Determine deflection limitations. For most applications, the allowable deflection of the
282 Sandwich construction
Facing failure Initial failure may occur in either compression or tension face. caused by insufficient panel thickness, facing thickness, or facing strength.
Of =
M t, hb
~
I
Tensile failure in facing
Transverse shear failure
V hb
= -
Caused by insufficient core strength or panel thickness.
Local crushing of core Caused by low core compression strength
u = -P
‘
A
General buckling Caused by insufficient panel thickness or insufficient core rigidity.
Shear crimping Sometimes occurs following, and as a consequence of, general buckling. Caused by low core shear modulus, or low adhesive shear strength.
Face wrinkling Facing buckles as a ’plate on an elastic foundation‘. It may buckle inward or outward, depending on relative strengths of core in compression and adhesive in flatwise tension.
PCr= tcGc
Adhesive bond
1/2
Core comDression
EftC
Intracell buckling (dimpling) Applicable to cellular cores only. Occurs with every thin facings and large core cells. This effect may cause failure by propergating across adjacent cells thus inducing face wrinkling.
Faces buckle into core cell
Fig. 12.22 Modes of failure in sandwich structures. Sandwich structures must be designed to resist these modes of failure. Failures may occur which combine more than one of the modes shown.
Structural analysis for specific cases 283 Maximum shear force V
Beam type
I
Simple s t m . .o r t
1
1
Uniform load
1-
Shear deflection constant
I f
Both ends fixed
Uniform load
Simple support
Center load
E
I I
I
IP
E
P -
A Both ends fixed
Center load
8
384
I
1
1 -
5 __
8
2
1
4
“ I + 192
1 1 -
192
1 4
I
I Cantilever
P = ql
I
I
1
I
I
I
Uniform load
lrrrrrrr
1
Cantilever
Triangular load
I
1
One end simply . _ supported .. one end fixed Uniform load
PL 8
I
Fig. 12.23 Loaded beam chart, where P = total load (per unit width), L = span, 0,= facing stress, t, = skin thickness, h = centroid distance, zcs= core stress, shear, tcc = core stress, compressive,A = (1 -/,L)~facing property, E , = modulus of elasticity of facings, Gc = modulus of elasticity of core in shear, Ism,= moment of inertia, sandwich, s = cell size, Ec = modulus of elasticity of core in compressions, F S = factor of safety, T = total sandwich thickness (note that P must be determined for a beam unit width). If deflections are critical, actual deflections should be verified by tests.
284 Sandwich construction the shear component should be calculated structure is usually limited to L/360. In and the core selection may be influenced by some cases, greater deflections may be the shear modulus needed. used, or, as in the case of snow skis, very much greater deflections may be a normal 9. Face wrinkling and intracell dimpling. With thin skins, a local failure of the skin in buckpart of the function of the structure. ling may be encountered. A check on the 4. Select skin material. Skin considerations afcrit will determine whether this may be a include the weight target, possible abuse design consideration. and local (denting) loads, corrosion or decorative constraints and costs. Select 10. Other considerations. Often, honeycomb panels are supported on more than two standard thicknesses and make the initial sides. If the ratio of length to width is calculation as outlined below. The facing greater than 3: 1, the calculations using the thickness directly affects both the skin stress shorter span and designing as a unit beam and the deflection. are quite adequate. The formulas in Roark16 5. Calculate first approximation. After the first are useful where the shear deflection may sandwich thickness, h, is determined, be ignored, using the following formulas. another selection of t, or E , may be made to arrive at more desirable or practical values t,h2b bt3 -~ of h. Most sandwich structures in ordinary 'sandw,ich 2 '. I solid =-- 12 usage are in the thickness range of So, for plate calculations: 1.5-150 mm (0.06-6.00in). 6. Select skin thickness. Keep in mind that mateIsolid = 6t,h2 rials such as fiberglass cloth and aluminum are available in specific, standard thick- Use of these formulas for deflections may give nesses. After the skin thickness for lower values than actually experienced, since deflection is selected, it should be checked the shear deflection may be important. Table for stress. The formula for 6, is used and a 88 of RoarkI6 gives some approximate multipliers to use for plates when supported as factor of safety determined. 7 Select core. Calculate the core shear stress, sc*. noted. Note that the core strength is not the same in the L and W directions. Refine the selection, 12.6.4 SIMPLE FORMULAS including considerations of material compatibility, cell sizes and types. Determine the Bending stress in facings: corrections needed to account for the effects M of thickness on strength, as shown in Fig. a, = th 12.5. Check the factor of safety using the calculated stress and the corrected allowable where M is determined by Fig. 12.23. stress. Other considerationsinclude crushing Core shear stress: and compression strengths, modulus in shear, weight and costs. For rolling wheel = -V loadings, the crushing strength and the skin cs h thickness are often the most important conwhere V is from Fig. 12.23. siderations. 8. Check def7ection. For many applications, the Deflection: calculation of the expected deflection may ~~1x32 K~PL omit the shear deflection portion. With a A=+E, t,h2 hGc very small deflection limitation, with a very thick sandwich, or with a very short span, ,
~
,
Structural analysis for specific cases 285 ( K , and K, from Fig. 12.23). For most beams, the second term is relatively small, but should be checked if deflection is critical or span is short.
material is to be woven roving, polyester and core to be KI'-3/8-60(25). Load, P': P'
Moment of inertia, sandwich:
Kb, Ks, M, V from Fig. 12.23:
~
Kb = 0.013, K, = 0.125, M = 8, V= 2.
-[I
2E, t,
,
ufcnt. = 1
Skin, t,: Try t = 0.090 in
Face wrinkling:
A, E , for fiberglass use:
A = 0.98, E , = 1.85 x lo6.
E,tc Factor of safety: FS =
120/144 = 0.833 psi
Span, L: L = 8 x 12 = 96 in
t,h2b I,, = 2 Face dimpling:
=
Calculate h:
Allowable or typical stress calculated stress
A=
tPE, -
12.6.5 SAMPLE PROBLEM: ANALYSIS OF FLAT RECTANGULAR SANDWICH BEAMS
Design a flat roof panel for a bus stop. Use a snow load of 120 lb/ft2. Use a simple panel with a simply supported span of 0.203 m (8 ft). Deflection is to be limited to L/270 and the factor of safety is to be greater than 2.0. Skin
K,P'L~~P
0.013 x 0.833 x 964x 2 x O.98ll2 0.090 x 96/270 x 1.85 x lo6
h = 5.518 (round out to 5.5 = panel total thickness, h = 5.41) Try thicker skin t, = 0.150 h = 0.013 x 0.833 x 96*x 2 x O.98ll2 0.150 x 96/270 x 1.85 x lo6 = 4.274
Use 4.00 overall thickness, h = 3.850. Since either construction is practical check out the skin and core stresses: P'L2 0.833 ~ 9 6 = ~ 1971 psi Uf = 0.090 x 5.41 x 8 t$Mc ~
Uf =
0.833 962 0.150 x 3.85 x 8
=
1662 psi
te that the skin stresses are quite close; therefore, the factors of safety would be similar. Fig. 12.24 Schematic diagram of a flat sandwich
panel.
FS
=
~
38 000 = 19;FS = 38000 = 23 1971 1662 ~
286 Sandwich construcfion For core: 'rs
-
P'L - 0.833 x 96 = 739 psi 5.41 x 2
or,
12.6.6 ANALYSIS OF FLAT RECTANGULAR SANDWICH COLUMNS: COLUMN DESIGN EXAMPLE L=Bfi
/
t
(20.32 cm)
0'833 96 = 10.38 psi 3.85 x 2
Note that the core stresses are quite low and there is not much difference in the stresses for the two thicknesses chosen. For KP-3/8-60(25), W shear strength = 60 psi. From Fig.12.5: thickness factor = 0.42, W shear modulus = 5800.
+P
Facings: Tempered hardboard,
W, shear, corrected = 60 x 0.42 = 25 psi
p = 0.99
FS = 25/ 10.4 = 2.4.
E , = 0.65 x lo6 psi
The use of KP-3/8-60(25) with a factor of safety of 2.4 could be marginal, which may vary from lot to lot of material. The other properties, compression strength and density are acceptable. Note that if the core is oriented to utilize the L shear properties, KP-1/2-80(11), with rCs= 70 x 0.42 = 29.4 might be satisfactory.
Core: Urethane foam, 6 lb/ft3, tc = 3 in From Table 12.1:
Lcs = 90 psi
Calculate deflection: For 5.50 T,A =
L r - c = 170 psi KbP'L42 ~
t,h2E,
+
KSP'L2 Gr = 1500 psi
~
G'h
-~0.013 x
0.833 x 964x 2 0.090 x 5.412x 1.85 x lo6 x 0.833 x 962 + - 0.125 5800 x 5.41
Check facing yielding:
Pp = 2t,oY= 2(0.25)(3600)= 1800 lb/in Check general buckling:
E , t,h2 Pcr =
= 0.377 + 0.032 = 0.0409 in
Note that the added shear deformation is only 9% of the total deflection. 0.013 x 0.833 x 96j x 2 For 4.00 T, A = 0.150 x 3.852x 1.85 x lo6
r l
L'+
(where D =
~
2(0.99)
0.125 x 0.833 x 96: 5800 x 3.85 PPV=
= 0.447 + 0.042 = 0.489 in
X'D n2D t,G'
D = 0.65 x lo6 x (0.25)(3.25)' =
+
= 3600 psi
866 872 lb-in/in of width n2(866872) (96)2 +
~ ~ ( 8 872) 66 3.0(1500)
Manufacturing sandwich stsuctuses 287 = 112 lb/in or 1352 lb/ft
temperature (both pressure and temperature in the precise amounts, at the precise time Check shear crimping: required for cure of the adhesive being used); and the provision for tooling and fixtures to PCr= tcGc= (3.00)(1500) hold the assembly in the desired shape and keep all the details in their proper positions = 3214 l b in/in of width during cure. Many different ways of providing these conditions are currently used, from vacCheck dimpling and wrinkling: Since facings are relatively thick and continu- uum bags or simple presses to autoclaves and ously supported by foam core, dimpling or unit tools, where volume and complexity can justify them. Most of the equipment is similar wrinkling will probably not occur. to equipment used in producing bonded structures or reinforced plastic parts where no 12.6.7 DESIGN CONDITIONS sandwich structure is involved. However, In-depth treatments for the design conditions bonding of sandwich structures is nearly listed in Table 12.7 can be found in MIL- always performed at lower pressures than is HDBK-23I, available from the US Government the bonding of structures which do not have a Printing Office. low density core and tooling is sometimes lower in cost as a result. Aside from the need for lower maximum pressure, there is little 12.7 MANUFACTURING SANDWICH noticeable difference between a sandwich STRUCTURES bonding facility and one which only handles The manufacture of sandwich structures non-sandwich bonding. A few suggestions can be offered to aid in requires three conditions to be met: the application of pressure; the application of living with the problems of sandwich bonding.
Table 12.7
MIL-HDBK-23 * CHAPTER
Subject ~~
_ _ ~
~~~
Wrinkling of sandwich facings under edgewise load Dimpling of sandwich facings under edgewise load Design of flat, rectangular sandwich panels under edgewise compression load Design of flat, rectangular sandwich panels under edgewise shear load Design of flat, rectangular sandwich panels under edgewise bending moment Design of flat, rectangular sandwich panels under combined loads Design of flat sandwich panels under uniformly distributed normal load Design of sandwich cylinders under external radial pressure Design of sandwich cylinders under torsion Design of sandwich cylinders under axial compression or bending Design of sandwich cylinders under combined loads Design of sandwich strips under torsion load Design of flat circular sandwich panels loaded at an insert
~~~
7 8 9 10 11 12 13 19 20
* M1L-HDBK-23 is revised from time to time, with new chapters sometimes added and older material updated. A check with the Plastics Technical Evaluation Center, US Army Armament Research and Development Command, Dover, New Jersey, USA, can verify that you are in possession of the most recent revision.
288 Sandzuick construction
1. Make sure the core is properly sized to fit the space it is intended to occupy. If it has been stretched a little, to make the distance from one edge member to the opposite one, it will probably shrink back as the cure cycles starts, leaving mysterious voids next to an edge member. If it is undersize in thickness at an edge, the adjoining edge member or fitting will hold the facing away from the core and result in an unbonded area. 2. If a honeycomb core is being used, remember that the adhesive between the core and the faces will end up much thinner than the same adhesive between the edges or solid inserts and the facings. For this reason, it is common to require the core to be as much as 0.25 mm ( 0.010 in) thicker than adjoining solid parts in the same assembly. 3. The elevated temperatures which most core-to-facing adhesives require for curing are often inaccurately measured. A good point to remember is that only the adhesive being cured can give you the cure temperature you are trying to measure. Some shops insert thermocouples directly into the bond line to determine temperature and then leave the thermocouple permanently in the part after cure is completed. 4. Most adhesives flow at an early point in the cure cycle. At this time, the bond lines will change in thickness by substantial amounts. The tooling employed to establish the shape of the part and hold details in place must also allow the details to move into their final cured position. Simple examples are a hot platen press, in which the platens close on the sandwich as the bond lines grow thinner, or an autoclave, in which a flexible bag follows the details as the adhesive flows, continuously transmitting the autoclave pressure to the shrinking assembly. Keep in mind that most adhesives are very weak and crack-prone as they go through the gel point. 5. Inserts or heavy members being cured as a part of a very light assembly will heat up
much more slowly, resulting in warpage problems upon cool-down. Warpage on very light parts can also be caused by one side cooling down too fast as a result of having one side removed from the still-hot tooling, while the other side continues to stay at the temperature of the tool. Also one side, next to the bag may be heating faster or to a higher temperature than the opposite face, which is in contact with a massive and still cold tool. Slower heat-up rates or better heat distribution in the tool design will help prevent these problems. 6. Be sure to provide a route for the escape of trapped air and gases from a totally enclosed part while it is being cured. This is particularly important in parts which are vacuum bagged to a female tool and cured in an autoclave. A coarse cloth 'breather' should be enclosed inside the bag to prevent the bag from sealing off portions of the assembly as pressure is being applied. Critical or expensive assemblies should have several vacuum lines attached at different points of the bag, with each monitored separately by a pressure recorder. 7. Caul plates should be carefully matched to the job they are expected to perform. These tooling aids are often used to cover the top of an assembly containing several different pieces of core, inserts, edges, etc., so that a thin skin will not push each detail to the minimum bond line thickness and result in an uneven outer surface. When the caul plate is moderately stiffer than the top skin, the bonding pressure is transmitted more to the thicker inserts and less to the undersized inserts, allowing all of the details to 'float' in the adhesive before cure, resulting in optimum relative placement of all the internal details in the sandwich. If the caul plate is extremely stiff or thick, this effect is changed to one of simply bridging over the most oversized details and the danger of producing voids or unbonded areas over the thinner details is substantially increased. Generally, the caul plate should
Manufact uring sandwich structures 289 not be more than two or three times the thickness of the sandwich facing material. Where thicker caul plates are used, the dimensional control over the size of detail parts in the assembly must be correspondingly better. The advantage of using such a thick caul plate derives from the ability to make both sides of a sandwich part have the smooth appearance usually associated only with the 'tool side'. 8. Make sure that core, pre-cured or rigid edges, inserts, skins and other relatively unyielding details assembled in the lay-up have close enough dimensional control to allow adhesives or resins to achieve the target strengths. In simple bonded assemblies, a tolerance of + 0.1 mm (+ 0.005 in) is necessary, while assemblies having multiple layers of prepreg or many layers of thin metal doublers can sometimes be successfully produced with much less demanding dimensional control. 12.7.1 CORE SHAPING
When core materials must be cut, trimmed, carved, or shaped, many special purpose tools are available. Sawing is the most common machining method, using either conventional blade tooth patterns, or, for some trimming operations, a special 'honeycomb band', in which the blade appears to be running backward, with the teeth sharpened on the back side, so that each tooth acts as a slicing knife blade. A different type of saw is also used as a mandrel-mounted router bit. Such tools, shown in Fig. 12.25, are very common where sculpturing of honeycomb or foam is to be accomplished. Router speeds vary from 1000-30 000 rpm for blade diameters of 1.8-10cm (0.754 in). Roll forming can be accomplished on metal cores, as shown in Fig. 12.26, while non-metal cores must usually be heat formed' In either forming can be much easier if an inherently formable cell configuration, such as that shown in Fig. 12.8 view H, is used.
CA t
.-
Fig. 12.25 Honeycomb carving bits employing a slitting saw 0.254 mm (0.010 in) thick x 12.5 teeth per cm (32 teeth per in), 50.8 mm (2 in) in diameter at the cutting edge. Turning at 12 000 to 30 000 rpm, these tools leave a smooth, burr-free surface on nearly any core material. The coarse teeth on the tool in the foreground are for the purpose of breaking up and removing the excessive amounts of core in cut depths of 5.08-50.8 mm (0.2-2.0 in).
Fig- 12-26Metal honeycomb may be roll-formed using Ordinary The surface usuallY must be protected during the operation by inclusion of a loose sheet of thin sheet metal between the core and the outer forming rolls. The tool being used is a 'Farham Roll', co-ody used in sheet metal shops.
290
Sandwich construction REFERENCES
Fig. 12.27 Nose radome core assembly, assembled by edge-bonding together several post-formed sections of glass fabric-phenolic honeycomb. Nomex core may also be formed in this manner.
I
Fig. 12.28 Effect of roll-forming on aluminum honeycomb. The core on the left has been roll-formed in sheet metal forming rolls, while the piece on the right has not been pre-formed at all. Note the anticlastic, or ‘saddle shape’, which the unformed piece assumes when forced to a cylindrical form.
1. MIL-HDBK-23, US Government Printing Office, Washington, DC*. 2. MIL-HDBK-17, US Government Printing Office, Washington, DC. 3. MIL-HDBK-5, US Government Printing Office, Washington, DC. 4. MIL-A-132, US Government Printing Office, Washington, DC. 5. MIL-A-25463, US Government Printing Office, Washington, DC. 6. MIL-STD-401, US Government Printing Office, Washington, DC. 7. Adhesive Bonded Aerosvace Structures Standard Repair Handbook, US’ Government Printing Office, Washington, DC, 8, Hexcel Corporation, TSB-120. 9. Hexcel Corporation, TSB-123. 10. Hexcel Corporation, TSB-124. 11. Alcore, TR-il2. 12. American Cyanamid, Handbook of Adhesives. 13. Baltek Corporation, Baltek Catalog. 14. Plantema, Frederic J., Sandwich Construction, John Wiley & Sons, New York. 1966. 15. American Plywood Association, Plywood Design Specifica tion. 16. Roark, R.J., Formulas for Stress and Strain, McGraw-Hill, New York, NY., 5th edn, 1975. 17. Timoshenko, S., Woinowsky-Krieger, S., Theoy ofPlates and Shells, McGraw-Hill, New York, NY. 2nd edn, 1959. *Publicationsof the US Government may be updated and revised from time to time. Be sure you have the most recent edition. This can be checked by contacting the Plastics Technical Evaluation Center, US Army Armament Research and Development Command, Dover, New Jersey. The publication MIL-HDBK-23 was abandoned some years ago. However, because the information it contained continues to be needed by designers of spacecraft structures, the entire publication will in future be included within MIL-HDBK-17.
METAL MATRIX COMPOSITES
13
V. I. Kostikov and V. S. Kilin
13.1 INTRODUCTION
Industrially developed countries are sucessfully producing materials and shapes based on metal matrix reinforced with carbon fibers (MMC). Metal matrix, ceramic and carbon composites reinforced with both discrete and continuous reinforcements have been described’-6. The present chapter will describe the results of investigations of the properties and preparation of aluminum/carbon composites MMC production deals with the solution of specific problems which involve the formation of useful structural components, with the attainment of their potential properties. The first problem is to fill the interfiber space of the carbon fibrous framework with matrix metal or alloy. Several molding methods are available. Optimization of the materials and the processes for obtaining usable MMC may further the use of MMCs in industry. When the MMC is reinforced with high molecular and high strength fibers the composite may be very brittle and can have thermal dynamic incompatibility with many metals used as matrices. MMC production technology is complicated and requires satisfaction of the following conditions, of which the most significant are as follows: 1. maintaining the reinforcing fibers’ strength; 2. ensuring a strong bond of fibers with matrices and between the matrix layers; Handbook of Composites. Edited by S.T. Peters. Published in 1998by Chapman & Hall, London. ISBN 0 412 54020 7
3. providing the correct fiber length, greater than the critical length; 4. even distribution of fibers in the matrix; 5. orientation of fibers in the direction of the applied load; 6. achieving the required shape and dimensions of the MMC; 7. obtaining a MMC strength reasonably near to theoretical.
Contemporary processes for obtaining an aluminum/carbon MMC can be divided into three main types: solid stage, liquid stage, and solution sedimentation. In the solid stage, the greatest effort has been spent in obtaining the MMC by hot extrusion. In this case, a foil, used as a matrix, is interspersed with carbon fibers to form a preform. The preform can then be subjected to optimum pressure, time and temperature in air, vacuum or inert gas medium. Considering metal bonding in the liquid stage, chemical bonds predominate; rarely are there mechanical bonds and there are no physical bonds7. Strong chemical bonds are possible because the atoms (or surface) of one substance come close enough to the surface of the other (1.5-3 A) to enter the zone of the surface atoms’ field-of-forces effectivity. When MMCs are produced by solid stage methods (diffusion welding, rolling, extrusion), it is practically impossible to provide full convergence of fiber and matrix surface due to uneven surface contours. Only application of extreme external forces makes the convergence possibleH.This pressure increase leads to brittle fiber breakage. Also, some
292 Metal inafvix cotnposites fibers inside the bundle may not be bonded takes place only after the surface catalysis by treatment in an oxidizing medium, sensitizing with the matrix because of a shadow effect. Due to the small diameters of carbon fibers and activation of the carbon fibers prior to and the short distances between them in the proper metallization. The method is complex tow (about several microns), poor bonding of because of the requirement for a strong, thin, matrix to the uncoated fibers located deep in well-adhering bond to the carbon fibers withthe bundle prevents development of their out compromising their mechanical or potential strength. Also, during the produc- physical properties. A typical treatment of cartion of composite material, MMCs based on bon fibers in an oxidizing medium of 65% metals such as Al, Mg, Ni, Ti, Cu, etc. and rein- solution of nitrous acid for 5 min does not lead forced with carbon fibers, the components to fiber strength loss. Ions and radical groups show high reactivity affecting composite per- that have affinity for metal are attached to the formance. Protective coatings (silicon, degraded surface of the carbon fibers. When silicon carbide is used as a barrier titanium, zirconium carbides, and nitrides) on the carbon fibers can prevent these undesir- layer, an oxidizing treatment of the carbon fibers is not necessary because the barrier able effects9J". Many MMCs can be made by soaking the coating itself is a good adhesive layer with a reinforcing fibers in the molten metal matrix. rough surface structure which bonds to fibers The problems can be solved by application of and nickel coating. In this case, preliminary a thin, strongly-bonded metal film that allows treatment for all types of carbon fiber tows the melted matrix material to completely wet and tapes includes only sensitizing with stanthe fiber. There are several ways to apply these nous chloride solution at 80°C for 10 min and coatings: stretching fibers through the melt; activation of palladium chloride at 80°C for 5 spraying the molten matrix metal; depositing min. During the sensitizing process, a metal film from the gaseous phase. However, hydrolized Sn2+ions which have high absorpapplication of a coating at constant thickness tion properties are strongly attracted to the on the carbon fibers in tows is not always fea- carbon fiber surface. During activation, pallasible because of fiber mutual screening. Fiber dium chloride is reduced to metal by the coating by a chemical method may be more ionic tin bonds formed at sensitization. Washing of fibers with water follows. The effective. activated fiber surface is then dried at 60-70°C for 15-20 min. 13.2 CARBON FIBER COATING METHODS Nickel coatings are applied from a solution Chemical methods are used to deposit thin of nickel chloride, 50 g 1-l, sodium hypophoscoatings and improve wetting on filaments in phite, 20 g I-l, ammonium chloride, and 50 g 1-', carbon tows and tapes . Because of low depo- trisodium citrate at 80°C at pH 8-9. A lustrous sition rates, chemical methods cannot be used nickel coating is obtained which varies in for deposition of thick layers of matrix. There thckness from 0.05-2pm at holding time of are two widely-used methods of deposition. 0 . 1 4 min. An estimate of coating continuity, and indirect data on coating adhesion to carbon fiber by means of a scanning electron 13.2.1 CHEMICAL PRECIPITATION microscopy, has shown that the coatings are (ELECTROLESS DEPOSITION) applied evenly and that the nickel penetrates Chemical deposition of nickel and copper on the microrecesses of the fiber surface, fully carbon tows and tapes with various textile replicating it and filling the grooves and irregstructures is based on the reduction of metal ularities. In carbon tows every individual fiber ions from a water solution. Nickel deposition is coated with nickel.
Carbon fiber coating methods 293 The textile structure of a braided fabric becomes more complicated. Coating thickness discontinuities may be considerable because of reduced clearances between fibers in zones difficult to access and stagnation and poor solution exchange take place. In these zones, the solution is fully depleted, while in other zones, the nickel plating may continue. The addition of 0.01 g1-1 lead sulfide to the sohtion slows the process of electroless nickel plating and eases the conditions of the formation of uniform coating thickness. Reducing the number of individual fibers in carbon tows to 100 and the number of ends in tows also help to provide better uniformity of coating thickness. When 300-fiber, 6 pm equivalent diameter, twisted tows of bean shaped cross section were coated, the coating thickness on the fibers was uneven and varied from 0.2-0.6 pm up to 0.5-1.0 pm at nickel gain in weight from 0.5 to 1.0 g. (Editor's note: Carbon fiber tows are not generally twisted and are all single end in the USA and other countries.) The thickness of nickel coatings, used for subsequent impregnation of carbon tows and tapes with aluminum melts, is usually about 0.5-1.0 pm. The nickel coating obtained by chemical deposition from solutions does not diminish the fiber strength. The chemical deposition of copper is accomplished at 20°C in alkaline solution that contains sodium sulphate, formalin, caustic soda, potassium sodium tartrate, and diethyl sodium carbonate. At a holding time of 3-6 min, the coating thickness reaches 0.1-0.4 pm. Before application of copper coatings, the fibers are subjected to the same preliminary treatment as the nickel coating. The more rapid electrolytic method of deposition should be a more efficient technique to obtain thick nickel or copper coatings. However, the electrolytic method may be used only for building up of the coating layer on previously electrolessly coated carbon fibers. Otherwise, the fiber tows will be coated with a crusty layer from the outside.
-
13.2.2 THERMAL DECOMPOSITION OF CARBONYLS
Coatings of refractory metals applied to carbon tows and tapes by thermal decomposition of volatile carbonyls in the gaseous phase have the following advantages: low temperature of decomposition of carbonyls of some commercially important metals (Cr, Mo, W, Rhs etc.); high reaction velocity; uniformity of coating thickness on each individual carbon fiber; ease of adjustment of structures and properties of coatings as a result of change of the deposition conditions. The thermal decomposition of the metal carbony is carried out as follows: Me(CO)n= Me + nCO In addition to the main reaction, side processes that cause the presence of admixtures of metal carbides in the coating can take place. As the temperature is increased, the possibility of obtaining pure metal layers without admixtures of free and fixed carbon occurs11. To obtain pure metals with a minimum content of admixtures, the process should be run at the highest possible temperature and sufficiently high vacuum. Some characteristics of coatings by precipitation from the gaseous phase are shown in Table 13.1. The type of substrate heating is usually determined by temperature requirements, and may be performed by conduction, induction, radio frequency or infra-red radiation. During the vapor phase of carbonyl decomposition there are a number of subsequent processes: vaporization (sublimation) of carbonyls, heat transfer, chemical process in the gaseous phase and substrate, adsorption-desorption on the substrate, formation of nuclei and growth of coating crystals12.Successively changing the carbonyl vapor feed velocity (growth of the coating layer from 0.3 to 6-8 pm h-l) and the temperature of substrate heating from 300 to 900°C can result in denser coatings, which will act as barriers excluding direct interaction of the matrix and the fiber, with a more
294 Metal matrix composites Table 13.1 Conditions for production of metal coatings by thermal decomposition of carbonyls from the gas phase
Heating temperature, "C
Metal coat ing ~....
cu Au V Cr Mo W Mn Te Re Fe Ni CO co Ru Pt Pt
-
_ _ ~
~
of carbonyl
of base
20 20 20 40 50 70 70 20 70 25 20 20 20 20 100 20
250400 120-150 70-100 350-700 450-700 450-700 110-300 60-70 400-600 100-350 100-250 180-220 180-200 200-300 500-600 210-220
~~~~~~
~~
At the initial deposition stages, thin smooth films reproducing the fiber surface are formed. At this point, one of the most important factors is the selection of an optimal coating thickness. This depends substantially on intended use of the material. Extremely thin coatings may be entirely soluble in the matrix and incapable of preventing counterdiffusion of the reacting components and thus may fail to ensure the barrier effect. Thicker coatings reduce the carbon fibers tensile strength, as shown below. The estimated strength of carbon fibers coated with silicon, titanium and zirconium carbides as a function of the corresponding element content is shown in Figs. 13.1 and 13.2. The element content is assumed to be equivalent to the carbide coating thickness, provided the coating is uniform. The strength change of carbon fibers with a silicon carbide coating derived from the substrate carbon is of particular interest [see Fig 13.1, coating deposition reaction (l)].The strength of carbon fibers (initial strength 2000 MPa and thickness 25 nm) at first increases at a coating thickness of 2-5 nm.
porous upper layer, providing the capillary effect for wetting and improving the fiber impregnation. 13.3 APPLICATION OF COATING IN GAS PHASE
A conventional method for the effective protection of carbon fibers to avoid reaction with metals is chemical vapor deposition (CVD) or refractory coatings. For example, when carbon fiber materials in the form of VMN grade tow and a carbon tape 'Kulon' are used as a substrate for the coatings, the coating is deposited from an appropriate gas mixture by one of the following reactions:
0
2
4
6
8
10
Si Content, Wt. 0;o
(1) MeCl,+ C (2) MeC1, + CH, (3) 2MeC1, + N, + 4H,
-+ -+
-+
MeC +2C1, MeC + 4HC1 2MeN + 8HC1
The coating thickness is controlled by the deposition time.
Fig. 13.1 Variation in strength of CF silicon carbide coating relative to thickness and composition of steam-and-gas mixture: (1) deposition reaction: SiC1, + CH, + Sic + 4HC1; (2) 3 deposition reaction: SiC1, + C -+ Sic + 2C1,.
Application of coating in gas phase 295
Fig. 13.2 Relation between tensile strength of titanium carbide and zirconium carbide coated CF and coating thickness.
This increase may be a result of healing of fiber "E surface defects. Then the strength of carbon E > fibers sharply decreases with increasing coaty" 50 ing thickness. In the case of carbides obtained r;w by the reaction scheme (2), the carbon fiber f ?? strength is stabilized at a certain level when a 5 30 definite thickness of the coating is attained. L% S a, The Young's modulus of carbon fibers with cn n a barrier is a structurally less sensitive factor; it 0 4 8 12 16 changes quite insignificantly and is usually Si, Zr, Wt. % somewhat higher than that of the fiber, irre"E spective of the initial coating thickness and E type. The shear strength of the carbon fiber 40 Y surface increases, apparently by surface defect &w healing and the strong bonds between the carf ?? bon fibers and the coating (Fig. 13.3)based on 8 20 the results of torsion testing. L% S a To determine the barrier properties, the 0) 0 compatibility of carbon fibers having a refrac0 2 4 6 8 tory coating with a metal matrix (AI, Ni) was Ti,Wt. % investigated by an internal friction method9,12 using suitable microcomposites in the form of Fig. 13.3 Relation between torsional strength of carbon fibers with a double coating. The nickel coated CF and coating thickness.
-
296 Metal matrix composites coating was deposited on the carbon fiber sur- not subject to noticeable structural changes face by CVD from an aqueous nickel salt connected with interactions of the carbon solution, and samples of the aluminum matrix fibers with nickel or the coatings. For carbon microcomposite were cut from plates obtained fibers with and without silicon carbide or zirby vacuum hot molding or vacuum aluminum conium nitride coatings, the maximum sputtering. Internal friction is the logarithmic internal friction background is observed at 600°C which indicates the structural changes decrement divided by x: to the carbon fibers. X-ray analysis showed the presence of nickel carbide. For the same car(13.1) bon fibers having no contact with nickel, the curve was similar to a curve for internal fricwhere A, and A,+I are the amplitudes of the tion change for fibers with titanium and zirconium carbide coatings. first and the nth vibration. The advantages of carbide coatings over Compatibility of carbon fiber and barrier coatings based on silicon, titanium, and zirco- nitride with thicknesses less than 300 nm were nium carbides, and titanium and zirconium shown when studying the compatibility of the nitrides with nickel additions (coating thick- coated carbon fibers with an aluminum matrix ness 50-100nm at up to 1200°C) is shown in (Fig. 13.5). Direct investigations confirming the barrier Fig. 13.4. It can be seen that up to 11OO"C, the internal friction background is negligible in coating efficiency have been conducted with C-Sic-Ni-A1, the case of titanium and zirconium carbides. the systems C-TIC-AI, This means that in the absence of pronounced C-Sic-Si-A1. The effect of carbide coatings on maximum and minimum, the carbon fibers are the carbon fibers-aluminum interaction was
1
280
ZrC Tic
0
200
400
600
800
1000
1200
Temperature, "C
Fig. 13.4 Relation between internal friction background of barrier coated CF and temperature in contact with nickel.
200
400 600 Temperature, "C
800
Fig. 13.5 Relation between internal friction background of barrier coated CF and temperature in contact with aluminium.
Application of coating in gas phase 297 investigated with nickel and silicon coatings. The degree of aluminum-carbon interaction was estimated in terms of quantity of the obtained aluminum carbide, using the aluminum carbide-water reaction. The results of determining the aluminum carbide content on carbon fibers with various coatings, and their tensile strength after annealing (the aluminum matrix is etched away) are shown in Table 13.2. These results suggest a strong aluminum-carbon fiber interaction. The quantity of aluminum carbide is sharply increased with annealing temperature and exposure time. However, a Sic coating with a thickness of only 28 nm decreases the aluminum carbide formation by a factor of 1/5. Doubling the Sic coating thickness inhibits the reaction resulting in the aluminum carbide formation by an order of magnitude. Subsequent increases of the coating thickness affect the aluminum carbide formation at a lesser degree.
Deposition of nickel and aluminum coatings on a Sic barrier layer substantially affects the carbon fibers-aluminum interaction. These coatings, dissolving in the aluminum melt, ensure impregnation of the carbon fiber tows. However, in the presence of nickel, the residual aluminum content on the fibers after annealing at 720°C is increased. This increase obviously is associated with a nickel-silicon carbide interaction that results in opening the fiber surface. However, the surface layer increases the Sic barrier properties. The quantity of aluminum carbide measured after annealing under the same conditions is decreased down to the level which can be obtained for Tic coating. This is associated with the fact that silicon dissolving in aluminum saturates it and suppresses the following reaction: (4)
Sic +A1
+
(Sic) +Al,C,
To summarize, based on the estimated strength of carbon fibers with barrier coatings and their compatibility with aluminum and
Table 13.2 The effect of coating thickness on the aluminium carbide content in the microcomposite A1-C and CF strength after annealing
Coating appearance and thickness
AI&
Temperature,
"C
content, rng m-=
Tensile strength,, GPa
Time, min. -~
Without coating Sic, 28 nm SIC, 56 nm
Sic, 97 nm Sic, 56 nm Sic, 69 nm Tic, 53 nm
+ Ni, 200 nm + Si, 230 nm
670 720 770 670 720 770 670 720 770 670 720 770 720 720 720
Initial CF (without coating) strength shown in brackets
5
10
20
10.5 22.2 28.3 2.2 4.9 6.2 0.6 0.8 1.0 0.5 0.7 0.7 1.2 0.4 0.3
17.9 44.8 47.0 2.8 5.8 10.9 0.9 0.8 1.4 0.7 0.8 1.4 2.7 0.5 0.5
35.1 50.9 54.5 4.3 6.5 17.7 1.1 1.7 2.6 1.0 2.0 2.7 3.1 0.8 0.6
1.43 (2.80) 1.96 (2.90) 2.34 (2.53) 2.28 (2.43)
298 Metal matrix composites
nickel, it can be concluded that titanium and zirconium carbides are preferred for a nickel matrix, and that silicon and titanium carbides are preferred for an aluminum matrix with the coating thickness of no more than 100 nm.
shapes, by rolling in dies and extrusion. The hot extrusion, and consolidating in autoclaves are used for manufacturing cylindrical cases and tubes with longitudinal and circumferential reinforcement.
13.4 TYPES OF COMPOSITE STRUCTURES
13.5 CONTINUOUS CASTING
Typical cross sections of MMC composite structures are shown in Fig. 13.6. Many types of continuous length composite structure can only be produced by continuous casting. In some cases, it requires a preliminary application of a barrier coating on the fibers by means of chemical deposition from solution or vapor-gaseous mixtures. It is convenient to prepare a unidirectional tape for composite structures by plasma spraying of matrix alloys on rows of monofibers or taped fibrous prepegs. In some cases, use is made of a variation of a hot extrusion treatment compression method to obtain a taped foil composite structure reinforced by one or several rows of fibers and infiltrated tows, as well as of plasma sprayed or infiltrated tapes. Sheets and plates are obtained by rolling in a die and extrusion. The simplest sections can be produced by bending of rolled strips, and more complicated
Preformed composite MMC with constant cross section can be manufactured by continuous casting or drawing of fibers through melt. The principal advantages of continuous casting are continuity, reduced time of contact of the fibers with molten metals and minimum unstable period. This method also results in a high production system, efficiency and reduced capital investment. The basic expenditures are for control equipment since the process only requires a common melting furnace and a foundry crucible, molding fixture, a magazine of spools with fibers and a system for transportation of fibers and handling products. During operation the melting furnace must be filled with argon or evacuated. As a result, the cost of the MMC produced by continuous casting is close to the cost of the initial fibers and matrix.
@
4 1
2
3
4
Fig. 13.6 Typical sections of semifinished items of composite materials; (1)filament with metal coatings and multifilament braids infilitrated with metals; (2) bars reinforced with braids or filaments; ( 3 ) bands with single-layer or multi-layer reinforcement; (4) pipes and cylindrical housings with longitudinal reinforcement; (5) sections; (6) bands and plates.
Infiltration under pressure 299 The continuous casting method has a number of advantages from a metallurgical standpoint since the composite is prepared without formation of typical oxide inclusions which are common in plasma spraying and hot molding. Most investigators have used vertical techniques of continuous casting, but there are versions of horizontal drawing of the fiber through the melt. Before being submerged into the melt, each fiber is detached from the other fibers to ensure intimate contact with the molten metal. Then the fibers pass through the drawing die that determines the cross section of the preform. The fibers are drawn with a rate necessary for solidification of the matrix to fill up the inner space of the composite preform. For example, in production of cast boron aluminum cable, the drawing rate could reach 900 m/h, but cannot exceed 450 m/h due to fiber defects. The continuous casting process can consist of two stages: the first stage involves the production of the composite preform with a small cross section, essentially a wire, bar or band; in the second stage, the previously produced composite preform is integrated into a more complicated structure with a larger cross section. For example, in the course of production of a boron-aluminum MMC, the most popular form of the initial composite structure is a bar of a circular section, containing 16-19 fibers. In the second stage, the amount of matrix added into the MMC can be very small to keep a high volume fraction of reinforcing fibers in the final product. The two stage process is used for the production of such profiles as rods, angle bars, T-beams, I-beams, and lenses. The composite resulting from cast lamina may be obtained in one, two or more steps. To form profiles, particular attention must be focused on the first stage since it determines the basic physicomechanical properties, especially specific strength. Because of the organization of the liquid phase process by stages, the amount of manual labor can be reduced, the equipment simplified and the dead time due to fiber breaks shortened. Also,
operations with fibers and the procedure for testing of fiber volume fraction are shortened. The production of aluminum carbon fiber for MMC is complicated due to several factors associated with the carbon fibers. Their small diameter (6-10 nm), friability, high reactability and poor infiltration characteristics with aluminum melts at temperatures below 1030°C are all contributory factors. The latter two are most important. The danger of chemical degradation and reduction of fiber strength due to graphitization and formation of carbides increases with greater contact time of the fiber with the melt and with higher temperatures. 13.6 INFILTRATION UNDER PRESSURE
lnfiltration of fibrous preforms makes it possible to obtain intermediate preforms and composite structures of limited size. Infiltration is accomplished by various means: by heating the mold containing the fibers and matrix material in the form of a foil, powder or coating on the fibers up to above the melting temperature of the matrix material with the subsequent shaping of the mold; by submerging the fibers into the melt or filling molds with the melt in vacuum, protective atmosphere or air; by vacuum suction of the melt; by feeding the melt into the mold under pressure; and by centrifugal casting. In all types of infiltration, the common denominator is the presence of the mold with the fixed fibers. Because the process is long compared with continuous casting, the fibers should have high thermal stability in metal melts. Experiments on infiltration with uncoated carbon fibers and with nickel plating by suction or by filling with melted aluminum or magnesium in vacuum have been unsuccessful due to the failure of the fiber to be infiltrated at the melt temperatures without extensive carbide formation. Common drawbacks of vacuum infiltration of carbon fibers with aluminum or magnesium alloys are incomplete infiltration because of nonuniform
300 Metal matrix composites penetration of the matrix material into the interfiber spaces, nonuniform package configuration and touching and burning of the fibers where there is no protective coating. These drawbacks are especially evident in infiltration of multifilament braids and bands. Uneven distribution of fibers results in defects in the form of isolated groups of filaments not separated from each other by matrix, thus serving as stress concentrators. The result is that with an increase in volume fraction of carbon fibers to more than 30-35%, which does not increase the strength of the composite material but starts to reduce it, due to the greater number of groups of filaments not infiltrated with the matrix. To improve infiltration of the carbon fibers with aluminum melts with metal coatings (especially nickel plating), conditions must be chosen to ensure a high rate of flow and to considerably shorten the duration of time of contact of the coating with the melt13.The relationship between the capillary and excessive pressure and the infiltration angle and volume fraction of carbon fiber in the molten aluminum matrix on the other, taken at a temperature range 925-1083"C, shows that the melt penetrates the braids of the fibers only at a pressure less than 1.1x lo5 Pa. Experimental testing shows that, under appropriate pressure, the rate of infiltration of fibers with aluminum melt with an 1%(mass) titanium additive moves at 0.5 m/s. The technology of infiltration of carbon fibers and fabrics with melts using aluminum has been advanced with application of external pressure. However, the best tensile strength values of the carbon/aluminum MMC with V , = 4040% are not in excess of 700-800 MPa, whereas the higher values (1000-1200 MPa) would be expected from rule of mixtures. To eliminate the strength drawback, it is necessary to determine the nature of barrier ceramic and process metal coatings which ensure protection (compared with the nickel plating) and allow effective infiltration of
braids of carbon fibers with aluminum melts at lower temperatures and higher pressures. It has been reported that a two layer coating produced the best results'l; the first layer of silicon carbide protects and the second layer of such transition metals as chromium, molybdenum, tantalum or tungsten which are properly infiltrated with aluminum improves infiltration. They also have a lower affinity for carbon compared to silicon. 13.7 ROLLING IN VACUUM
Nondeforming process methods, including ionic and plasma spraying and spontaneous infiltration make it possible to obtain the high physical and mechanical properties of carbon-aluminum MMC. However, these methods are characterized by low productivity and high labor input. The formation of the structure and properties of carbon aluminum that has been vacuum rolled within the solidification range of the matrix aluminummagnesium alloy has also been studied. Rolling of the initial billets of carbon aluminum was performed on a vacuum rolling mill with residual pressure Presid = 6.5 x lo-' Pa. Powder alloys were used as the matrix, the carbon band, grade 'Kulon' with two-layer coating (silicon carbide and free silicon) was used as reinforcement.The temperature conditions of the rolling were characterized as follows:
where T is the rolling temperature within the solidificationrange of the matrix alloy; TIlqand Tsolare the temperatures of the liquidus and solidus of the matrix alloy. The amount of reduction of porosity in the initial porous billets of carbon aluminum was determined by considering the maximum spread and extraction of the part of the matrix alloy from the billet in compliance with the following formula:
Rolling in vacuum 301 = 1- [ m c / p c+ (mm- m'J
/~,1/(1.15 1, bo ho)
(13.5) (13.3)
where
mc and mm are the masses of the carbon fibers and matrix alloy in the billet; p, and p, are the densities of the carbon fibers and matrix alloy; 1, bo h, are the initial values of the length, width, and thickness of the billets; rdrnis the mass of the matrix metal extruded from the billet in rolling.
where Vf is the volume fraction of fibers in the material; V;" is the volume fraction of the fibers which are not surrounded with matrix alloy; VF is the volume fraction of the fibers within the volume of the fibrous material limited by the contours of the fiber braid.
When the metal completely fills the fiber interSince maximum properties of the composite stices and the fibers are completely well material can only be reached with a 100% fill- distributed in the matrix K and R are equal to 1. Figures 13.7and 13.8show the relationship ing of the interfiber spaces inside the braids of between factors K and R and the modes of liqthe carbon fibers and the fibers must be evenly uid phase rolling for carbon-aluminum. The distributed within the matrix, the filling (K) processes of filling up of the interstices and and structural evenness ( R ) factors of the the formation of a uniform structure are fibrous material are: highly dependent on the temperature and Vf - vy K= (13.4) degree of each reduction in rolling. When there is a low amount of liquid phase, the Vf
Y c
c
0 al L
al
Q
0
0.2
0.4
0.6
0.8
1 .o
Temperature Coefficient, q~
Fig. 13.7 Infusion percentage, K of braids CF MCM AI-C for various rolling modes: (1)& = 1.0 EP; (2) E 0.75 E ~ (3) ; E = 0.5 E ~ growth ; = 6.65 x lo-' Pa.
=
302 Metal matrix composites
0
0.2
0.4
0.6
1.o
0.8
Temperature Coefficient, y~
Fig. 13.8 Uniformity, R of structure of MCM A1-C for various rolling modes: (1)E (3) e = 0.5 eo; growth = 6.65 X lo-' Pa.
=
1.0 eP; (2) F = 0.75 Ep;
braids of the carbon fiber are sealed without CR = N - N R (13.6) filling up of the interfiber spaces. Thus, ?,bn > N 0.6 due to an avalanche-like decreasing level of hydrostatic pressure at the deformation cv = N-NV (13.7) point and efficiency of filling up of the braids N is reduced. The latter is associated with physical changes in the billet during deformation where: from the solid body with limited inclusions of CR is the relative number of the kinked the solid phase. Increasing the rate of rolling fibers; and increasing the number of the steps to the Cv is the number of the broken fibers; same total reduction improves the conditions N is the number of the fibers in the specimen; for forming a uniform structure in carbon NR is the number of the fibers that have lost aluminum, and the factors K and R approach their original cross-sectional shape unity. (kinked); The highest quality composite must be Nv is the number of broken fibers. sealed. During rolling, within the solidification range of the matrix alloy, the number of During these experiments, it was found that shrinkage pores obtained in the composite the probability of fibers breaking during solid requires seal rolling of the matrix in the solid phase rolling can be reduced by increasing phase. However, the probability of breakage number of steps before sealing. and kinking of fragile carbon fibers also When these process requirements are satisincreases. The degree of breakage and kinking fied by rolling under vacuum, carbon of fibers has been estimated using relative fac- aluminum composite with the following tors during microstructural research of the structural characteristics can be obtained: rolling specimens: fiber braids are filled to 95-97'/0; structural ~
Rolling in vacuum 303 uniformity is 78-94%; and the number of defective fibers is 4-10%. High quality rolled carbon aluminum is most dependent on the percentage of defective fibers. In a 45% Vf composite, the properties in the direction of reinforcement are as follows: tensile strength, 825 MPa; bending strength, 1300 MPa; and the modulus of elasticity is 200GPa. A two layer protective coating on the carbon fiber, with a small amount of aluminum carbide (- 0.1./0), retains the transverse strength of the reinforced composite (a, = 55 MPa). Application of a protective chromium coating on the carbon aluminum by condensation ion bombardment also increases its resistance to corrosion. However, carbon aluminum is somewhat unsatisfactory in economic terms because of its inherent problems with soldering, welding and mechanical connection with other materials. These drawbacks can be eliminated by using combined composites of the following systems: boron-aluminum - carbon-ahminum (high compressive strength, good resistance to erosion); titanium - carbon-aluminum, dispersively reinforced aluminum (filling agent: Sic, A1,CJ carbon-aluminum (high hardness, transverse strength). There are specific criteria for production of such composites using the method of rolling along the fibers. It is advisable to use some of these criteria to optimize the manufacturing process15. The basic requirements for carbonaluminum are: sealing of the porous billets during extrusion; 100% filling of the interfiber spaces with matrix alloy; provide for compatibility; and prevent component oxidation. Research considering the influence of rolling manufacturing parameters on meeting criteria requirements makes it easier to select an optimum technology for production of carbon aluminum. The main feature of this technology is that rolling is performed over 10-12 steps at a temperature corresponding to the median value of the solidification range of the matrix alloy (- 40-50 vol. YOof the liquid phase). After solid phase additional sealing,
the rolling must be accomplished at a vacuum of 6.65 x lo-' Pa, and compacting of the initial powder-like billets of the carbon aluminum must be performed with rigid, plastically undeformed casings. However, this technology is unsuitable for production of combined composites. For example, during production of Ti-A1-C by this technique, intimate contact of the layers of titanium with carbon-aluminum is not realized; whereas, during rolling of fibrous material A1-B-C, an intensive interaction of the boron fibers with the molten aluminum takes place. Thus, in the production of combined composites, the criterion requirements are necessary with at least an additional requirement for intimate connection of the separate hybrid components. To produce a diverse combination, such as MCM AIB-MCM A1-C, the number of requirements doubles. However, research shows the requirements can be met by non-traditional manufacturing approaches. The combination Ti-A1-C is an example of the product of such an approach. The bimetal titanium-aluminum composite can be successfully produced by rolling at temperatures of up to 500°C (932"F), with high degrees of deformation. The requirement for preservation of fiber continuity and the elimination of the carbon-aluminum reaction decreases the probability of formation of an intimate connection between titanium and aluminum layers during production of the combined composite. Also, during compaction of carbon-aluminum within the solidification range of the matrix alloy, the probability of formation of friable intermetallic layers between titanium and aluminum is very high. Thus, the application of liquid phase methods (infiltration, soldering, and rolling above the alloy solidus temperature) makes it possible to obtain a high quality combined composite. For the combined composite, the following manufacturing methods were used. The pack billet consisted of layers of the matrix alloy, titanium foil, and carbon bands and was heated to a temperature
304 Metal matrix composites exceeding the liquidus temperature in the vacuum chamber of the rolling mill. Preferred evaporation of magnesium from the aluminum alloy with subsequent deposition on the contact surface of titanium, as well as simultaneous infiltration of the carbon bands took place. After treatment, the billet was cooled down to the carbon-aluminum compacting temperature and was vacuum rolled. Deposition of the protective magnesium coat does not form chemical compounds with titanium. The infiltration of the carbon framework with subsequent welding-soldering of titanium with matrix alloy during rolling ensures that bonding bridges are formed during the liquid phase high temperature contact of components at the moment the infiltration is completed. Plate and layer reinforced carbon aluminum with high resistance to corrosion, with transverse strength exceeding that of the aluminum by 1.5-2.0 times, has been obtained by this technology. m s composite can also be successfully attached to titanium structures by soldering or welding. The strength of the titanium to matrix alloy bond is higher than the strength of the bond between the carbon fibers and the matrix (attempts to peel the titanium result in fracture between the aluminum and the carbon). To obtain the combined composite A 1 - K , the following ground rules have been established: compacting of plasma porous semifinished boron aluminum items should be performed in the solid phase at temperatures exceeding 560°C (1040"F)15; and carbon-aluminum should be compacted in the middle of the solidification range of the matrix alloy. An alloy of A1 with 10% Mg with a solidus temperature of 510°C (950°F)was used as a matrix for carbon-aluminum, but a higher temperature alloy was used as a matrix in the boron-aluminum. The assembled pack billets of boron-aluminum and carbon-aluminum unidirectional lamina were rolled at a temperature of 540-550°C (1004-1022°F) for 9-10 steps (the compacting mode for carbon-aluminum). With tlus procedure, filling of the interfiber space in
the braids of the carbon fibers, as well as the infiltration of the porous monolayer of boron aluminum with molten melt, but without direct contact with the boron fiber, takes place. After cooling below the solidus point of the carbonaluminum additional reverse rolling results in preferred sealing of the boron-aluminum. Analysis of the microstructure and mechanical properties of the obtained material shows proper weldability and filling of separate monolayers, with a minimum number of the broken fibers. This combined composite shows better mechanical properties than carbon-aluminum. Thus, for a one step filling technique, the combined composite has 10-15% greater tensile strength than the carbon- aluminum composite and the transverse strength was increased by a factor of 2 (20% B). Combined fibrous dispersively-reinforced composites AIM-Sic and Al-C-Al,C, have been obtained by careful selection of the matrix alloy or by using an alternative manufacturing processes. For these composites, adherence to the criterion requirements is relatively simple. It is possible to produce sealed monobands from the dispersively reinforced composites (AI-Sic, Al-C-AI,C,) with subsequent compacting of the composite using the techniques developed for compacting carbon-aluminum. Data developed on compatibility of carbon fibers with barrier (protective) coatings and aluminum and nickel matrixes, as well as data on strength as a function of a barrier coating thicknesses, have facilitated the development of a number of MMCs. Properties for some of these MMCs are presented in Table 13.3. These materials were made with 'Kulon' carbon fibers with Russian production barrier coatings. Data on these fibrous materials both coated and uncoated are shown in Table 13.4. Figures 13.9 and 13.10 show articles made of aluminum-carbon and nickel-carbon composites. Although MMC properties are excellent, higher strength carbon fibers will permit further increases of both specific and absolute strength of the newly created MMC.
Rolling in vacuum 305 Table 13.3 Some properties of carbon reinforced metal-matrix composites Property Filler Tensile strength M,,, GPa Elastic modulus, min, GPa Coating
Matrix AL-9
AMG-6
Kulon 2.5 40.0 Tic
Kulon 2.5 40.0 Sic
__
Ni
cu
VMN-4 2.0 25.0 Ti(Zr)C
VMN-4 1.5 23.0 -
Metal-matrix composite Density, p, kg m-3 Tensile strength, MPa at 20°C at 400°C at 800°C Bending strength, MPa Compressive strength, MPa Elastic modulus, E, GPa Specific modulus, E/p, m2s-* x lo6
2000-2300
2200-2300
5200-5400
5200-5500
900-1000 800-900
700-900
450-500
400-500
-
-
-
-
-
900-1050 800-1000 270-300 135
800-1100 500-700 240-270 109
400-500 700-900
600-700
-
-
140-180 26
120-170 23
Table 13.4 Some properties of carbon fibers without coating and with barrier coating Properties Density, g cm” lb in-3 Filament diameter, pm Elongation at break, YO Carbon content, wt. YO Coating Coating thickness, nm Specific surface area, m2g-’ Longitudinal CTE X 10“ at 21°C (70°F) Tensile strength, MPa (psi) Tensile modulus, GPa (psi) Typical Strand Properties: Threads/strand Filaments/ thread Twist, tpm (tpi) Twist / strand Filaments/strand
VMN4
VMN4SiC
VMN4TiC
Kulon-Sic
1.72 0.062 6.5 1.10 298 w/c 51.0 2.25 (1.2)
1.72 0.062 6.5 1.10 295 Sic 60-80 50.5 2.25 (1.2)
1.72 0.062 6.5 1.10 295 Tic 50-80 50.5 2.25 (1.2)
1.91 0.069 6.5 1.10 294 Sic 60-80 10.5 2.25 (1.2)
22000 (290 000) 2230 (33 x 106)
22300 (334 000) 2230 (33 x 106)
22300 (334 000) 2230 (33 x 106)
22500 (363 000) 2400 (57 x 106)
24 300 100 (2.54)
24 300 100 (2.54)
24 300 100 (2.54)
-
-
-
7200
7200
7200
24 300 100 (2.54) 2 600
-
306 Metal matrix composites
Fig. 13.9 Various types of shaped articles made of aluminocarbon composites material: (1) shell; (2) tube; (3) gas-turbine blade; (4) T-profile.
REFERENCES 1. Lee, S.M. International Encyclopediu of Composites. New York: VCH, pp. 187-358. 2. Fridlyander, I.N. Metal Matrix Composites (Soviet Advanced Composites Technology Series). London, Chapman & Hall. 3. Metal matrix composites. The next generation of high-performance materials. Emerging Technologies, No 20 Technical Insight Inc., Englwood/Fort Lee, NE, May 1986, p. 190. 4. Bunsell A., Gorg, M. Les matriaux composites a matrice metallique an Japon. Industrie Ceramique. 1988, No 830,642. 5. Zimick D.G., Koik B.M. Design of thermally stable graphite/aluminum tubular structures for space applications. SAMPE Q., 1990, 21(2), 11. 6. Strength optimization of C/A1 metal matrix composites produced from prepregs. Ph.D. Thesis. Masson J.-J. DTR-FB-92-39, ETN-9394871.1992. p. 110. 7. Ivanov, V.S. (ed.). Aluminium and Magnesium Alloys, Reinforced Firbres. M. Science, 1974, p. 202. 8. Fukuda, Sh., Matsubara, T., Tokao, Yo. A fabrication process for fiber-reinforced composites by using a wire explosion spraying method. Carbon fiber reinforced aluminum preform sheets. Yosha, 1990,27(1),3540. 9. Dergunova V.S. et al. Protective coatings for carbon fibrous materials. In: Anticorrosive Coatings. L. Nauka, 1983. pp. 164-168.
Fig. 13.10 Ring made of the composite material nickel-carbon fiber.
10. Clement, J.R., Kock, H.J., Wu, K.T., Spenser, H.G. Interfacial modification in metal matrix composites by sol-gel process. Muter. Manut Process, 1990, 5(1), 17-33. 11. Syrkin, V.G. Chemistry and Technology of Carbonyl Materials. M. Chemistry, 1973, p. 185. 12. Moskalenko, A.G. et al. Kinetics of carbon fiber and aluminium contact interaction. Mekhanika Polymerov. 1977, No 4,4243. 13. Varenkov, A.N., Kostikov, V.I. Vliyuanie cmachivania grafitovykh volokon na ysloviya izgotovleniya materialov metodom vakuumnoy propitki. (Influence of wetting graphite fibers on conditions of manufacturing of composite materials by using vacuum impregnation method). In Volokriistye i dispersno-uprochnenye kompozitsionnye materialy (Fibrous and DisperselyReinforced Composite Materials). M. Science, 1976, pp.25-28. 14. Shorshorov, M. Kh. et al. Methods of improvement of infiltration of carbon fibers with aluminium. Adhesion Melts Soldering Mater. 1976, No 1,86-89. 15. Shorshorov, M. Kh. et al. Electronic and microscopic researching of carbonyl metal coatings carbon fibers for reinforcement of composite materials. Phys. Chem. Metal Treatment. 1976, No 4,141-143.
CERAMIC COMPOSITES
14
M.F. Amateau
14.1 INTRODUCTION
The mechanical characteristics of ceramic composites are profoundly influenced by the Ceramics generally are compounds of metallic mechanisms of strengthening and toughening. or non-metallic elements and other non-metals Successful design of ceramic composites such as oxygen, nitrogen, carbon and boron. requires an understanding of the role of the Compared to metals these compounds have higher melting temperatures, higher Young’s constituents, including reinforcement, matrix moduli and hardness, lower densities and and interphase, on the these mechanisms. In lower electrical and thermal conductivities. addition to the properties and form of the conEngineered ceramics are used in thermal and stituents, the characteristics of the processing method can also have a major impact on structural applications requiring high temperature resistance, high hardness and chemical mechanical and physical properties. inertness. Applications that exploit the thermal structural properties of ceramic 14.2 CONSTITUENTMATERIALS commonly include cutting tool inserts, wear 14.2.1 REINFORCEMENTS resistant components, ballistic armor, heat exchangers, burner tubes, prosthetics, dental Fibers used in ceramic matrix composites fall implants, heat engine components and ther- into three general categories based on their mal barrier coatings. The dominating diameter: monofilaments, textile fibers and characteristics of ceramics that limit these and whiskers. In addition, reinforcements in the other engineering applications are their lack of form of particulates and platelets are used in plastic behavior at room temperatures and ceramic composite designs. The strengthening their low tolerance to flaws, i.e. low fracture role of the reinforcements in ceramic compostoughness, that lead to catastrophic failure. ites is significantly different from that in Reinforcing ceramics with particles, polymer and even metal matrix composites. In whiskers, platelets, discontinuous fibers and metal and polymer composites the reinforcecontinuous fibers significantly improves their ments contribute directly to the increase of strength, toughness and apparent ductility. strength and stiffness by carrying a significant Composite design can also be used to tailor portion of the load. In ceramic composites the other important properties such as high tem- reinforcements usually increase the strength perature strength and thermal shock indirectly by increasing toughness of the resistance, wear resistance and low friction, matrix. The load carrying capabilities of the thermal and electric conductivity and thermo- reinforcements are of secondary interest at elastic proper ties. most. This places special importance on the whiskers as reinforcements for ceramic comHandbook of Composites. Edited by S.T. Peters. Published posites. in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
308 Ceramic composites Whisker reinforcements Whiskers are a very fine filamentary form of a material. They are usually 1 pm or less in diameter and up to 200 pm in length. They are nearly perfect single crystals with strength properties approaching the theoretical strength of the material. Figure 14.1 is a scanning electron microscope image of silicon carbide whiskers. As a reinforcement for ceramic matrix composites it is their size and aspect ratio that is most significant to their strengthening effect. The whiskers that have recently been commercially available are listed in Table 14.1. Silicon carbide whiskers obtained from different sources have unique surface compositions
Fig. 14.1 Scanning electron microscope image of Sic whiskers.
that can profoundly influence their interfacial properties and hence effectiveness as a reinforcement in a ceramic matrix. Surface layers of the whiskers can contain crystalline SiO, or amorphous S i 4 4 (socalled 'black g1ass')l. Surface compositions can also be modified by chemical treatments such as HF washing. Ceramic textile fiber reinforcements Although significant improvement of mechanical properties of ceramic materials can be achieved by whisker reinforcement, truly mechanically robust and reliable structures can only be approached with continuous reinforcements. Continuous ceramic fibers of sufficient strength and optimum stiffness and diameter are now commercially available. They are capable of being filament wound, woven or braided into textile forms. They are compositional variations of three systems: A1,03, S i 4 and Si-N. Table 14.2 lists the composition and selected properties of fibers that are commercially or near commercially available. The usefulness of these fibers can be measured by their high temperature properties and their ability to be shaped into complex preforms for subsequent infiltration. A convenient measure of the ability to handle these fibers for processing is their critical bend radius, p,, i.e. the smallest radius that the fibers can be bent before they fracture. A simple beam strength
Table 14.1 Selected properties of commercially available whiskers', Diameter (mm)
Length (mm)
Young's modulus (GPa)
Coefficient of thermal expansion (10-6PC)
Density (g cm-j)
Advanced Composite 0.6 Materials Corp. Tokamax Toaki Carbon 0.3-0.6 Company XPW-2 Huber Company 0.05-1.0 UBE Industries 0.05-0.5 SNWB 4-7 Catapal XW Vista Chemical Company
10-80
600
4.5
3.2
50-200
400-700
5
3.2
5-100 5-100 40-100
600 370 400
4.5 2.5 7
3.2 3.2 4.0
Whisker Grade
Sic Sic Sic Si,N,
A1,0,
SC-9
Source
Constituent materials 309 Table 14.2 Commercially or near commercially available continuous ceramic fibers
Manufacturer
DuPont Sumitom0 3M 3M 3M 3M DuPont Nippon Carbon Ube Dow Coming Dow Corning Dow Coming Tonen
Designafion
Composition
FP
>99 a-Al,O, 85 A1,03, 15 SiO, 73 A1,03, 27 SiO,
Tensile strength OMPa)
Tensile Density modulus Cg cm-’) (GPa)
Diameter Critical bend (pm) vadius (mm)
>1400 1800-2600 2000
385 210-250 193
3.9 3.2 3.03
20 9-17 10-12
2.75 0.53 0.48
62 A1,0,, 14 B,O,, 15 SiO,
1750
154
2.7
11
0.48
70 A1,0,, 2 B,O,, 28 SiO,
2100
189
3.05
10-12
0.045
70 A1,0,, 2 B,O,, 28 SiO,
2275
224
3.05
10-12
0.049
A1,0,, 15-25 ZrO,
2100-2450
385
4.2
20
1.83
50 Si, 31 C, 10 0
2520-3290
182-210
2.55
10-20
0.36
MPS
Si, Ti, C, 0 69 Si, 30 C, 1 0
>2970 1050-2450
>200 175-210
2.4 2.65
8-10 10-15
0.27 0.83
MPDZ
47 Si, 30 C, 15 N, 8 0
1750-2450
175-210
2.3
10-15
0.50
HPZ
59 Si, 10 C, 28 N, 3 0
2100-2450
140-175
2.35
10
0.33
Silicon nitride
60 Si, 37 N, 30,0.4 C
2500
300
2.5
10
0.60
Altex Nextel 550 Nextel 312 Nextel 440 Nextel 480 FP 166 Nicalon Tyranno
analysis will show that the p,, can be calculated studies are critical to developing proper comby multiplying the failure strain of the fiber by posite fabrication methods since almost all the radius of the fiber. Thus high strength, low processing techniques require high temperaelastic modulus and fine diameters produce the ture processing. Both FP and PRD166 fibers most robust fibers. The first successful alumina can be heated to 1000°C before any serious based textile fiber was DuPont FP fiber. This reduction in strength occurs3.The boria modifiber has a very high elastic modulus but its low fied alumina-silica (Nextel 312) and mullite strain to failure results in very large critical compositions (Nextel 440 and Nextel 480) are bend radius which has limited its application. A also degraded with high temperature heat modification of the all alumina fiber by adding treatments as seen in Fig. 14.24.The high tempartially stabilized zirconia, the DuPont FP 166 perature degradation in strengths and elastic fiber, has reduced the critical bend radius by moduli of the boria modified alumina silica one third. However, only the silica modified bases fibers are summarized in Table 14.3. The alumina based fibers such as the 3M Nextel retention of properties at high temperatures fibers and the Sumitomo Altex fiber have criti- following the relative degree of crystallinity among the three grades with Nextel 312 being cal bend radii of about 0.5 mm. Numerous studies have been performed on the least crystalline and Nextel 480 being the these fibers to determine their strength and most crystalline. Strength and Young’s modulus at temperamodulus retention after heat treatment. These ture of the Sumitomo alumina silica fibers has
310 Ceramic composites
480
4 -*-(===t-A
1oOO-
alr hydrogen
L.
'"r=
vacuum
Fig. 14.2 The effect of heat treatment temperature on the strength of alumina and mullite fibers (440 and 4800)4. Table 14.3 High temperature degradation of boria modified alumina silica fibers4 Property
Comparative temperature ("C)
Benchmark
~.
Strength after heat treatment
Hot strength Hot elastic modulus Creep rate
50% of initial in air in hydrogen in vacuum 50% of initial 50% of initial 10-4/h at 69 MPa
been measured by Bunsel15u p to 1200°C. Both the strength and elastic modulus begin to degrade at 800°C while above 1000°C the properties drop significantly. During the last 20 years there have been significant advancements in the development
Nextel 3 12
Nextel 440
Nextel 480
>1300 1200 1200 1000 950 800
1400 1100 1350 1120 1000 1010
>1400 1200 1400 1200 1250 1190
of a silicon carbide based textile fiber for composite reinforcement. All commercially available fibers in this category contain oxygen but can also contain nitrogen and titanium. Nicalon fiber manufactured by Nippon Carbon Company and marketed in
Constituent materials 311 the USA by Dow Corning Corporation is by far the most commercially developed. X-ray diffraction analysis indicates that Nicalon consists of ultra fine p-Sic particles dispersed in a matrix of amorphous SiO, and free carbon6. Nicalon has excellent resistance to thermal degradation in argon and air exposure at temperatures to 1000°C for as long as 100 h7.The loss of tensile strength for Nicalon by exposure to temperatures to 1400°Cin both air and a r m n are nrewnted in FiP 14 7
- 3 -
0"
t
9. b
d
g 2 g!
r
n
-
Monofilament reinforcements Monofilament Sic and boron fibers are produced by chemical vapor deposition onto a fine substrate filament. For the case of Sic fibers the core is 37 pm amorphous carbon filament, while for boron a 13 pm tungsten wire has been commonly used. The principal advantage of monofilament reinforcements are their ability to tolerate some degree of surface reaction with the matrix during fabrication or high temperature service. These fibers can be infiltrated by a number of processing methods including powder sintering, powder hot pressing, plasma spraying and melt infiltration. These fibers are limited to structures with relatively simple shapes such as sheet, plates and large diameter cylinders because of their large critical bend radius. Table 14.4 lists the properties of these fibers.
Q,
Table 14.4 Properties of monofilament reinforcements
e l -
f
'Original fiber 0
I
I
I
I
I
I
I
I
.
Manufacturer Composition Tensile strength (MPa) Tensile modulus (GPa) Density (g ~ r n - ~ ) Diameter (pm) Critical bend radius (mm)
(4
0
0
Boron
SCS-6
Textron B 2.5 400 2.5 140 11
Textron Sic 4.3 427 3 140 7
14.2.2 MATRIX MATERIALS
0'O\14000C
'0
0 1
10
1d
1o3
Heat treatment time, t (h)
(b)
Fig. 14.3 Loss of strength of NicalonTMafter exposure to (a) argon and @) air at temperature to 14OO0C7.
The selection of matrix materials for ceramic composites is strongly influenced by thermal stability and processing considerations. The properties of matrix materials commonly used in ceramic composites are shown in Table 14.5. These include oxides, carbides, nitrides, borides and silicides. The first indication of the ability of a material to resist high temperature service is melting temperature. With the exception of glass ail these materials have melting temperatures above 1600°C. As the melting temperature increases the ease of processing decreases.
312 Ceramic composites Table 14.5 Properties of typical ceramic matrix materials ~~
Materials
Young’s modulus (GPn) -.
LAS Pyrex Alp, Mullite ZrO, PS ZrO, FS
TiO, Si,N, SN Si,N, RB Si,N, HP SiOz Sic Sn Sic HP B4C TiB, Tic TaC Be0
wc
Cr $4 Cr& BNL BNII
NbC
117 48 345 145 207 207 283 310 165 310 76 331 414 290 552 427 283 359 669 103 386 34 76 448
Poisson’s Modulus of rupture ratio (MPa) __
0.24 0.20 0.26 0.25 0.23 0.23 0.28 0.24 0.24 0.24 0.16 0.19 0.19 -
0.20 0.19 0.24 0.24 0.20 -
0.20 -
0.21
138 55 483 186 648 248 83 496 303 827 -
386 462 310 896 248 200 234 -
262
Thermal expansion
Fracture toughness (MPa m”?)
Density
2.42 0.08 3.52 2.20 8.46 2.75 2.53 5.60 3.41 5.60 0.77 4.94 4.94
2.61 2.23 3.97 3.30 5.75 5.56 4.25 3.18
5.76 3.24 8.64 5.76 7.92 13.5 9.36 3.06
-
-
-
6.92 -
-
3.85
-
-
76 110
-
-
-
-
(g
Me1t ing point (“C)
(10-6PC)
3.19 2.20 3.21 3.21 2.41 4.62 4.92 14.50 3.00 15.80 5.21 6.70 1.94 1.94 7.82
3.06 0.54 4.32 4.32 3.06 8.10 8.46 6.66 5.76 4.50 7.56 9.67 6.66 0.36 6.66
-
1252 2050 1850 2760 -
1849 1870 -
1870 1610 1980 1980 2350 2900 3140 3880 2530 2870 2435 1890 2982 2982 3499
Mechanical and chemical compatibility of the 14.3 PROCESSING METHODS matrix with the particular reinforcement ultiProcessing of ceramic composites can be permately determines whether a useful formed by solid, liquid, or gas phase composite can be made. For the case of processing of the matrix material to achieve whisker reinforced composites the chemical infiltration of the matrix around the reinforcreactions with matrix are particularly critical ing phase. The goals in processing ceramic since even minor reactivity can consume the composites are to achieve minimum porosity entire reinforcement. Large differences in the with a uniform dispersion of the constituents coefficients of thermal expansion between and controlled bonding between the reinforcreinforcement and matrix can lead to large ing phase and the matrix. residual stresses during the fabrication and ultimately result in serious degradation of mechanical strength. Small or optimum differ- 14.3.1 POWDER PROCESSING ences can be beneficial to mechanical Fundamental steps in processing ceramics properties by placing the weaker constituent composites from powdered constituents are: in compression or by inducing crack deflection 0 powder selection; between reinforcements. 0 powder characterization; 0 agglomerate reduction;
Processing methods 313 constituent mixing; green body fabrication; green dressing (machining and gate removal); binder removal; consolidation and densification; final dressing (burr removal); inspection.
To minimize voids and interfacial weakness and maximize the toughening effect of the reinforcing phase, a uniform finely dispersed mixture must be produced. Arranging the constituents to minimize free space between them is referred to as ’packing’. When the constituents are not effectively packed, subsequent densification becomes difficult, requiring higher pressing temperatures, presThe selection of constituent powders is the sures and duration. Both constituent shape first step in composite design and consideraand particle size difference can affect packing. tion must be given to chemical, mechanical Optimum packing occurs when the particle and thermoelastic compatibility between the size distribution contains 30 vol. % of very constituents as well as the desired final small particle and 70 vol. % of large particles’. mechanical and physical properties of the If uniform round fibers (or whiskers) are percomposite. In addition to the obvious problem fectly aligned in a closed packed array then of reaction between constituents, other incommatrix particles approximately 0.15 times the patibilities such as large differences in melting fiber diameter would theoretically fit in the temperature of matrix and reinforcement can interstices. preclude successful processing. Thermal Most ceramic powders can be comprised of expansion mismatch between constituents can a mixture of primary particles and agglomercause premature failure in them or at their ates. Agglomerates are primary particles interface. bonded by surface chemical forces, electrostaThe rational selection of constituents usutic forces or solid bridging. In order to produce ally requires knowledge of certain physical a finely dispersed, homogeneous mixture of characteristics such as particle size distribumatrix and reinforcement successfully, the tion, shape, specific surface area, bulk density, agglomerates must be reduced. A typical electrical charge, impurities, etc. The ultimate agglomerate (mullite in this case) is shown in aim of such constituent characterization is to Fig. 14.5. The agglomerate is 8-9 ym in diamepredict the final characteristics of the ceramic ter while the constituent primary particles composite, as shown schematically in range from 0.1 to 1.5 pm. To uniformly incorFig. 14.48. porate 0.5-1 ym diameter whisker or particles,
MILL TIME
LUBRICANTS
Fig. 14.4 Use of powder characterization in process control8.
314 Ceramic composites
Fig. 14.5 Typical agglomerate found in mullite powders.
the large agglomerates must be broken down by mechanical action or chemical treatments if the agglomeration is due to surface forces. If the whiskers are robust or some degree of whisker breakage can be tolerated, both agglomerate reduction and constituent mixing can be accomplished simultaneously by ball milling. Organic binders are usually mixed with the particle-whisker mixture for near-net-shape processing by a variety of cold forming operations including uniaxial pressing, cold isostatic pressing, tape casting, extrusion, compression molding and injection molding. The ceramic preform after cold consolidation is referred to as the 'green' form. The part in the green form can usually be machined without damage. In this state additional near-net-shape processing can be applied such as gate removal and machining. Final consolidation and densification is performed at high temperatures. Three of the most common methods are sintering, hot (unidirectional) pressing and hot isostatic pressing. For low fiber or whisker contents
(5 % or lower) sintering may produce satisfactory results. For complete densification of even low fiber volume fraction composites, sintering may impractical due to excessive temperatures and durations. For high fiber or whisker volume fractions, hot pressing and hot isostatic pressing are the only effective methods for densification. Table 14.6 shows the effect of hot pressing time and temperature at 31 MPa pressing pressure on the theoretical density of Sic whisker-Al,O, composites for various vol.% whiskers. Theoretical densities of over 99 vol.9'0 can be achieved in unreinforced A1,0, at 15OO0C, in lO%SiC whisker composites at 1650°C and in 2O%SiC whisker composite at 1800°C10. 14.3.2 LIQUID PROCESSING
When high temperatures and mechanical forces are used to consolidate composites from the powder, the optimum strength properties can be sacrificed. Reducing processing temperature, time and pressure can minimize damage to the reinforcements but fully dense
Processing methods 315 Table 14.6 Effect of processing parameters on the theoretical densities of Sic whisker-Al,O, composites pressed at 31 MPa pressure'" Vol. %
Pressing temperature
Densify
("C)
Pressing time (mid
Density
zuh iskers
(g cnz-?)
(% theoreticnl)
0 10 10 20 20 20 20
1500 1500 1650 1500 1650 1725 1800
60 35 60 120 60 60 25
3.95 3.78 3.89 3.68 3.72 3.78 3.81
99.1 96.7 99.5 96.1 97.1 98.6 99.5
or near-fully dense composites cannot generally be produced. Processing by infiltration with a molten matrix would be an ideal way of minimizing mechanical damage and still achieve fully dense structures. The melting temperatures of ceramics used as matrices in composites limits the general use of melt infiltration as a viable processing route. However, by careful tailoring of the matrix and the use of innovative in situ reaction techniques, melt infiltration has been successfully utilized to fabricate ceramic composites. Glass and glass-ceramic matrices have been successfully infiltrated in the liquid form into fiber preforms by 'matrix transfer molding'".The high temperatures required to achieve the appropriate fluidity of the matrix limjts the available fiber-matrix compositions to only those with low mutual reactivity. Other matrix materials such as CaSiO,, SrSiO, and
SrO~Al,O;SiO, were infiltrated into Sic powder preforms with and without Sic whiskers with resulting open porosity of about 1%12. Recently considerable attention has been applied to directed melt gas-metal reactions which produce ceramic matrix composites directly from the liquid metal13.14. Both metal oxide matrix and metal nitride matrix composites have been produced by this technique. Net shape composites can be processed at temperatures of the melting temperature of the metal. The commercial development of this processes is called the DIMOXTMprocess of the Lanxide Corporation. In this process oxidation or nitridation occurs on the surface of the molten metal forming a layer of solid ceramic. The layer thickens as the molten metal wicks up between the grains of the ceramic. A schematic representation of the process is shown in Fig. 14.614.The phenomenon is made
Reinforcement preform Reinforcement preform entrapped in solid reaction product
Fig. 14.6 Directional metal oxidation method for processing ceramic composites4.
316 Ceramic composites possible by dopants which modify the surface energy between the phases. For instance, if the grain boundary energy, yB, is less than twice the energy of the solid-liquid interface, ysI.and the surface energy of the solid-liquid interface, ysL is greater than surface energy of the solid-vapor interface, ysv, thickening of the oxide (or other reaction compound) layer does not take place. By reversing the relative values of the surface energies, (i e, yB > 2ys1. and ysr < ysv) as illustrated in Fig. 14.7the unstable grain boundary permits wicking of the liquid metal through the grain boundaries of the reaction product phase'-'. Application of this technology to ceramic matrix components is achieved by allowing the reinforcement preform to float above the liquid metal bath for infiltration of the molten metal. A growth barrier can surround the reinforcement preform to produce practical net or near net shape component^'^, 16.
I
Molten Metal
I
14.3.3 VAPOR PROCESSING
The infiltration of the reinforcing phase by a gas that decomposes to form the solid matrix phase is generally referred to as 'chemical vapor infiltration' (CVI).Various carbides, nitrides, oxides and boridesI7 as well as unreacted carbon'* have been deposited on silicon carbide-based yam fibers (e.g. Nicalon and Tyranno fibers), oxide based fibers (e.g. Nextel fibers) carbon yam fibers and Sic whiskers. Silicon carbide is one of the most commonly applied matrices using CVI. Methyltrichlorosilane is reacted with hydrogen on the surface of the fiber to deposit silicon carbide. A typical reaction for this process is19. CH,SiC1,
+ H, + Sic + 3HC1 +H,
This reaction can take place by conventional chemical vapor deposition (CVD) at temperatures of 1000-1400°C. Silane-hydrocarbon (SiH,-C,HJ mixtures can be used to deposit Sic at temperatures below 500°C by plasma assisted chemical vapor deposition (PACVD). Table 14.7 lists some of the more commercially important matrix materials that can be applied by conventional CVD. A more complete list of ceramic materials produced by both conventional CVD and PACVD can be found in reference (20). Table 14.7 Ceramic materials formed by CVI processes2'
Matrix ceraTnic ~
I
Molten Metal
I
Reaction temperature ("C)
~~~
Tic Sic B,C TiN Si& BN A1N *1,0,
Fig. 14.7 Mechanism of directed metal oxidation growth14.(a) no growth due to stable grain boundary; (b) oxide growth mechanism with unstable grain boundary.
Reactant gases
SiO, TiO, ZrO, TiB, WB
Tic,-CH,-H, CH,SiCl,-H, BC1,-CH,-H, TiC14-N2-H, SiCl,-NH,-H, BC1,-NH,-H, AlCl,-NH,-H, AlCl,-CO,-H, SiH-C0,-H, TiC1,-H20 ZrC1,-C0,-H2 TiC1,BC13-H, WC1,-BBr,-H,
900-1000 1000-1400 1200-1400 900-1000 1000-1400 1000-1400 800-1200 900-1100 200-600 800-1000 900-1200 800-1000 1400-1600
Design considerations 317 The main drawbacks for processing composites by CVI are the high processing time and costs. Since the deposition occurs most rapidly on the outer surfaces, the internal passages can be blocked off long before full densification is complete. It is usually necessary to interrupt the infiltration process to grind the surfaces in order to reopen the gas access to the fibers or preform in the center of the part. Residual porosity of 10-20% with less than 10% open porosity are typically obtainedz1. Two basic methods of CVI are isothermal processing and forced flow/thermal gradient processing. In isothermal processing the fiber preform is heated by radiation from the walls of the furnace (so called ’hot wall reactor’) or by inductively heating a carbon mandrel on which the preform is placed. In both cases the decomposing gases are allowed to diffuse through the fiber preform. In the forced flow/ thermal gradient method the reactant
gases are forced through the fiber preform retained in a graphite holder with a sharp thermal gradient maintained by water cooling. A schematic diagram of the forced flow/thermal gradient method is shown in Fig. 14.8. 14.4 DESIGN CONSIDERATIONS
The approach to designing ceramic matrix is constrained by the brittle nature for both the matrix materials and reinforcements used. Unlike polymer matrix composites and even metal matrix composites, the rationale for design of ceramic composites is to impart toughness in a structure that would have unacceptable toughness as a monolithic ceramic”. Design methods are unique to the form of the composite, depending on whether continuous unidirectional reinforcements, discontinuous reinforcements or multi-layer, multi-directional reinforcements are being considered. As a starting point in the design of continuous
Exhaust aas Heating element Perforated lid
---
Infiltrated preform Fibrous preform
Reactant gases Fig. 14.8 Forced flow thermal gradient method for CVI processingzo.
318 Ceramic composites strength over the range of reinforcement volume fractions will depend on the relative fracture strain, strength and stiffness of the constituents. The relative fiber and matrix stress-strain curves and strength prediction of a composite consisting of a high stiffness, high strength fiber in a lower stiffness, low strain to failure matrix is represented in Figs. 14.9(a) and 14.9(b),respectively. There are many such fiber-matrix combinations that have this relative behavior as the examination of Tables 14.2 and 14.5 will reveal. For instance if Nicalon is selected as the fiber then the selection of mullite, lithium alumino silicate (LAS) or Pyrex 14.4.1 DESIGN OF CONTINUOUS glass, for the matrix meets the requirement. UNIDIRECTIONAL REINFORCEMENT Selecting Sic monofilament produces this case COMPOSITES for almost all matrix materials listed in Table The Young’s modulus of unidirectional contin- 14.5 with the exception of TiB, and Tic. The uous fiber ceramic composites Ec is composite strength in such a system should satisfactorily predicted by rule-of-mixtures: increase at a rate predicted by the linear ruleof-mixtures based on the strength of the Ec = E , V , + Em V, matrix and the stress on the fiber at the fracture strain of the matrix, a;. At fiber fractions where E , and Em are the Young’s moduli of the to V,,,, failure of the matrix constitutes failure reinforcement and matrix respectively and V , of the composite. The behavior of the composand Vm are the respective volume fractions. ites with fiber fraction below Vcr,thave simple When there is a high bond strength between linear stress-strain behavior to failure. Above the fiber and matrix, prediction of composite this fiber fraction the matrix breaks before the
unidirectional ceramic composites, the rule-ofmixtures can be used to calculate elastic and thermoelastic properties of the composite. Strength properties of the ceramic matrix composites are poorly predicted by the rule-of-mixturessince flaw sensitivity and reinforcement-matrix bond strength are not addressed by these tecluuques. Rule-of-mixtures properties are less important in discontinuously reinforced ceramic composites since toughness is strongly controlled by the interfacial properties.
%
Strain
0
Vcrit
1
Fiber Fraction
(b) Fig. 14.9 Strength prediction for high stiffness, high strength fiber and a lower stiffness, low strain to failure matrix.
Design considerations 319 fibers. The fibers can retain the broken matrix X', is between the range given by in place before the fibers break at a higher load. Composite strength above Vcritdepends upon the fiber strength. A typical stress-strain curve for such a system above V,,, is shown where T is the interfacial shear stress and r is the fiber radius. The value of strain at the end of this process, E ~ isz3 ~ ~ ,
u)
u)
0
L
z V
Emu
E max
Strain Fig. 14.10 Stress-strain behavior for composite with high stiffness, high strength fiber and a lower stiffness, low strain to failure matrix.
schematically in Fig. 14.10. The elastic portion of the curve is followed by a serrated, constant stress portion induced by a matrix failure process. During this process the matrix continues to crack until the spacing between cracks,
The final linear rising portion is the curve is the elastic response of the fiber. Continuous fiber breakage and fiber pull-out can produce the pseudo-ductility of the final portion of the curve. There are many potential continuous fiber-matrix combinations in which the matrix has a higher elastic modulus than the fiber. For instance, matrix materials such as titanium diboride, titanium carbide, silicon carbide and alumina with most of the continuous fibers listed in Table 14.2 would have the relative stress-strain behavior shown in Fig. 14.11(a). The strength of predictions of such systems is shown schematically in Fig. 14.11@).In this case the predicted strength of the composite would decrease with increasing fiber fraction until a minimum fiber fraction, V,, is reached. This behavior is similar in appearance to the
sf"
(D
8
b
v)
Of
(b)
Fiber Fraction
Fig. 14.11 Strength prediction for a high strength fiber and a higher stiffness, low strain to failure matrix.
320 Ceramic composites case of a high modulus, low failure strain fiber transverse elastic moduli E, and E,, respecin a lower modulus, high failure strain matrix tively, composites with aligned short fibers can as is typical of many metal matrix composites. be made by using the Halpin-Tsai relation? However the cause of the minimum behavior is quite different. Below Vmh failure of the matrix still constitutes composite failure where the rule-of-mixtures strength is composed of the matrix ultimate strength, omuand the stress on the fiber at the matrix failure strain, a;. Unlike the case for the lower modulus matrix, the stress 0; is lower than the a,, increasing fiber fraction lowers the rule-ofmixtures strength. Above Vmi,the fracture of tL= 2 l / d , and tT = 2 the matrix no longer constitutes composite fracture as the fibers alone are able to carry the The coefficients of thermal expansion in the load after matrix failure. longitudinal and transverse directions, a, and The above description applies to composaT respectively, can be estimated fromz6 ites with a high fiber-matrix bond strength and neglects the effect of fracture surface a, = (a,E,V,+ amEmVm) / E, energy. These conditions are not typical in real aT = (1+ vf) a,V, + (1+ vm)amVm- aLvLT composites and the simple rule-of-mixtures predictions must be modified to account for where these effects. Aveston et aLZ4accounted for the VLT = vf + YmVm effect of the fiber-matrix bond strength on the matrix failure strain as follows: and vf and vm are the Poisson's ratio for fiber 6zTE,Vf 1/3 and matrix respectively. These calculations will E'f = EkVmrEc] usually overestimate the value of these properwhere r is the fracture specific fracture energy ties because of ineffective bonding between fiber and matrix and deviation from ideal fiber of the matrix.
v,
[
14.4.2 DESIGN OF DISCONTINUOUS REINFORCEMENT COMPOSITES
In polymer and metal matrix composites it is usually desirable to design the fiber lengths to exceed the critical length, 1, given by ufr/t to allow the fiber to carry its full load prior to fracture. In ceramic composites, fiber breakage is rarely the design goal. Instead the role of the fiber is to provide toughness by a combinations of fiber pull-out, crack deflection and crack bridging. Nevertheless, the designer may want to predict the elastic and thermoelastic properties of the discontinuous reinforcement composites' An estimate for longitudina1 and
Fig. 14.12 Microstructure of 20 vel. yo SiC-Al,O, composite fabricated by tape casting.
Design considerations
321
alignment. Figure 14.12 shows the in-plane laminate fracture theories must be employed. microstructure for a 20 vol% Sic whisker-aluExamples of material designs that can make mina composite fabricated by tape casting and use of laminated-composite concepts for hot pressing. This processing method promotes improved performance are illustrated in Fig. fiber alignment in the tape cast direction, how- 14.133z.The magnitude of the surface compresever there is still a considerable deviation from sive stress can be calculated from laminate the predominant fiber directionz7. theory. Figure 14.13(a) shows a laminate For randomly oriented fibers or whiskers the design intended to produce surface compreselastic modulus, E , of the composites can be sive stresses. In this design the layers toward estimated from the results of the Halpin-Tsai the mid-plane gradually increase in coefficient method using the empirical relation: of thermal expansion. The outer layers, con-
ET= (3/8)E, + ( 5 / 8 ) E , 14.4.3 DESIGN OF MULTILAYER, MULTIDIRECTIONAL REINFORCEMENT LAMINATES
The concept of a laminated composite is used effectively in the design of polymer-matrix composites to achieve the high degree of strength, elastic and thermoelastic tailoring. Polymer-matrix composites reinforced with either continuous or discontinuous fibers are fabricated by stacking layers with specific characteristics and orientations in a predetermined sequence to achieve desired mechanical or physical properties. As with polymer composites, the ceramic composite layer properties may be calculated using theoretical and semiempirical method^^*-^^ from the constituent properties such as the elastic modulus of the fibers and matrix respectively, E, and Em, the orientation factor, f, the volume fraction of the fiber, Vf, the fiber aspect ratio, l/d, and the coefficients of thermal expansion for the fibers and matrix, afand am.By selection of the sequence of layer orientations and compositions, various elastic, thermoelastic, strength, physical and chemical characteristics can be produced. Classical laminate plate the0ry3&~~ can be used to accurately predict the elastic and thermoelastic properties of laminated composites from the layer properties. The strength properties, on the other hand, cannot be readily determined by commonly used laminate failure criteria since fracture of these laminates is still strongly controlled by the presence of flaws. Modified
(a) Design with graded composition
I
(b) Design with toughening layers
Oxidation resistant layer Wear resistant layer High toughness core
(c) High temperature wear design
Fig. 14.13 Typical laminate design concepts for ceramic matrix composites. (a) with graded composition; @) with toughening layers; (c) for high temperature wear.
322 Ceramic composites taining increasing amounts of low-expansion whiskers generate compressive residual stresses as a result of the differential contraction during cooling after the high-temperature densification process. A major advantage of laminated-composite processing is that it provides the engineering flexibility to use innumerable material and property combinations that would be impossible with traditional methods involving thermal or chemical tempering. This concept also allows the use of non-equilibrium compositions for greater degree of stress profile variation. For instance, the depth and magnitude of the stress gradient can be independently controlled by selection of layer composition and properties. Maximizing the stress gradient by the introduction of a high-expansion material in the interior of the composite would be impossible by conventional chemical tempering but is quite feasible by lamination. Strengthening can also be achieved by rendering surface flaws ineffective through the introduction of a tougher ceramic layer below the surface (Fig. 14.1303)). This design mitigates surface damage in the outer layers by blunting the cracks when they reach the underlying toughened layer. This layer may contain whiskers, a toughened ceramic, or metallic particles. The use of a toughened ceramic layer as the outer layer would not be as effective since abrasion or impact could produce flaws through its entire depth, thus permitting the crack to propagate through the lower-toughness interior layers with minimum resistance. In addition to increased strength and toughness, high-temperature corrosion resistance can be designed into a composite material by using a corrosion-resistant layer on the exterior surface (Fig. 14.13(c)) and layers tailored for high-temperature strength in the interior. A similar concept may be employed for a material designed as a high-temperature heat exchanger by grading the interior layers for high thermal conductivity. Using composite laminate theory, a materials designer can tailor the grading to minimize the deleterious residual tensile
stresses that are likely in such a construction. Differences in elastic modulus and coefficients of thermal expansion for layers containing different volume fractions of reinforcing whiskers can be used to generate favorable residual stress patterns in fabricated laminates. The thermal stresses o,T,oy' and T~~ in each layer of the laminate at any position through the thickness, z, measured from the midplane, caused by the restraint of the neighboring layer can be determined by Hooke's
[:
]=
y:'
Alll A'l2 A' Af12A'22A':6 "16
Af26
{ q%v]
Properties 323
where t, is the thickness of the kth layer, Q, are the untransformed stiffness coefficients and al are the coefficients of the thermal expansion in the principal material directions. The thermal moments Mx, My and MXyare zero. The residual stresses, oLand oT,in the longitudinal and transverse directions respectively are
For this laminate geometry the residual shear stress, rLT= 0. The compressive residual stresses thus induced in the outer surface of the ceramic composite raise the fracture strength by that amounP3.
(a) Bridging
(b) Pull-Out
14.5 PROPERTIES
The principal objective in design of ceramic composites is to produce enhanced toughness and mechanical reliability. Various energy absorbing mechanisms are produced by the reinforcement depending on the relative thermal expansion coefficients, relative elastic moduli and interfacial bond strength between the reinforcement and matrix. In addition the size, shape, distribution and volume fraction of the reinforcement plays a strong role in controlling the effectiveness of the toughening.
(c) Deflection
14.5.1 MECHANISMS OF STRENGTHENING
The four principal mechanisms of toughening (crack bridging34,35, fiber pull-out, crack deflection and matrix microcracking) are shown schematically in Fig. 14.14. More than one of these mechanisms can be operative at the same time in a ceramic composite but there
(d) Microcracking
Fig. 14.14 Toughening mechanisms for ceramic matrix composites.
324 Ceramic composites is usually a dominant one depending on the resistance. A quantitative treatment of the constituent and interfacial properties. effects of crack deviation on toughness have Bridging, pull-out and deflection are most been provides by Faber and Evans3'. In certain composites the conditions can be effective where the fibers are generally aligned favorable to allow the stress field of the propnormal to the crack surface. In the crack bridging mechanism (Fig. agating crack to interact with the stress field 14.14(a))the fibers remain intact for some dis- around the reinforcements to produce local tance behind the crack front, thus restraining matrix cracking around the reinforcement. the crack opening displacement and reducing Maximum effectiveness of this mechanism the stress intensity at the crack tip. The energy requires a fine dispersion of many reinforceabsorbing processes include fiber fracture, ments as illustrated in Fig. 14.14(d). An fiber-matrix friction and elastic strain energy appropriate mismatch in thermal expansion of the fiber. Thus, strong fibers, high strain to coefficients between reinforcement and fracture fibers and strong fiber-matrix bond- matrix provides the local stress field around ing promote this mechanism. The toughening the reinforcement. An analysis of this mechaproduced by a uniform closure stress of the nism of toughening has been provided by bridging fibers has been estimated by Becher H ~ t c h i n s o n ~ ~ . Typical fracture in whisker reinforced et a1.35to be ceramic (SiC/Al,O,) shown in Fig. 14.15 has elements several mechanisms of toughening. dKc = (3fU[VfrEcG,/6(1 - v2)EfGi]l12 Whisker pull-out and crack defection are eviwhere Gm and Gi, are the strain energy release dent in this example.
"
between the fiber and the matrix. For fibers of critical length the increment in toughness is given as (see Reference 36, for example): AGc = V ,o f u r / 6 ~
Crack deflection mechanism (Fig. 14.14(c)) forces the crack to deviate out of the normal stress plane as it negotiates around the reinforcements. The driving force for this deviation is the residual stress distribution produced by the mismatch in thermal expansion between the fiber and matrix. Reinforcements with higher coefficients of thermal expansion than the matrix will cause the matrix to be in compression near the reinforcement. This state will tend to deflect a crack as it approaches the vicinity of the reinforcement which is a higher crack resistant area to a region of lower crack
Fig. 14.15 SEM photograph of fracture path in Sic whisker-Al,O, composite. 14.5.2 TYPICAL PROPERTIES
The selection of materials for component design requires accurate and reliable mechanical property data. Because of the large variations in processing characteristics and starting material forms, such data are generally sparse. The mechanical properties most signifi-
Properties 325 91 1 cantly affected by reinforcements in ceramics are fracture strength and fracture toughness. 8 Vaughn et ~ 1examined . ~ ~ the effect of processing temperature and reinforcement for alumina 7and Sic whisker/alumina composites. Table 14.8 shows that processing temperature and remforcement have little effect on elastic modulus, but a significant impact on strength, fracture toughness and work of fracture. The noncompliance with the rule of mixtures for Young’s modulus is evidence of the lack of bonding and hence load transfer between fiber and matrix. The toughening of various ceramic matrix materials with increasing Sic whisker content is shown in Fig. 14.1635. A compilation of fracture strength and frac0 0.1 0.2 0.3 0.4 ture toughness for Sic whisker (Silar-SC9) Whisker content (Volume fraction) /A1,0, composites is given in Table 14.92.The Fig. 14.16 Increment in fracture toughness of Sic fracture strength of the composites for the whisker composites with various matrices35.
Table 14.8 Mechanical properties of polycrystalline A1,0, and Sic whisker /A1,0, matrix composites3’ Alumina (1500°C) Young‘s 371 modulus (GPa) Fracture 456k40 strength (MPa) Fracture 3.3d.2 toughness (MPa m1/2) Work of 10 fracture om-’)
Alumina (1659°C)
(1900°C)
Composite (Silar SC-1)
380
375
375
393
385+18
2534
641+34
606+146
5.0~0.2
3.7d.1
8.7M.2
4.6d.2
39
67
21
20
Alumina
Composite (Tateho SCW-1-5)
Table 14.9 Room temperature strength and fracture toughness of Sic whisker (Silar-SC-9)/A120, composites (Adapted from Reference 2 ) Whisker content (Vol. %)
Fracture strength (MPa)
Fracture toughness (MPa mIn)
Xejerence
0
150 253 391 475 540 652 675 680 641 720 640 850
4.3 3.7 3.6 4.0 4.8 4.6 6.1 8.7 8.7 7.0 7.9 6.2
40 39 41 42 42 41 42 40 39 42 42 43
5 10 15 20 30
40
326 Ceramic composites same whisker content can vary fairly significantly among the different sources while the fracture toughness is relatively consistent. Above 30% whiskers some composites can exhibit either a strength or fracture toughness that actually decreases because of the decrease in homogeneity of the whisker distribution. The fracture strength and fracture toughness for Sic whisker reinforced mullite are shown in Figs. 14.17 and 14.18 respectivelp. These composites were fabricated by tape casting and the L-type designation indicates that the crack propagation direction is normal to the tape casting direction and the T-type designation indicates that the crack propagation direction is parallel to the tape casting direction. Greater strength and toughness are achieved for the case of fiber orientation predominantly normal to the whisker axis direction. The maximum strength and toughness for this system is at 40% whiskers. The effect of Sic (Silar S-9) whisker content on strength and toughness of Si,N, matrix composites is given in Table 14.1045.For this material the maximum toughness occurs at 30% wluskers while the strength maximum is 1000
I
I
I
I
I
I
8
20
30
40
50
Fig. 14.17 Fracture strength of Sic whisker-mullite composites with the crack propagating normal (Ltype) and parallel (T-type) to the predominant whisker directionM.
I
I
I
I
I
~
I
I
-
-
z 6Y
E 7
-
-
5-
-
-
2
26 '
Ib
io
' io ' ' 40 ' 5'0 ' SIC Whisker Fraction ( ~ 0 1 % )
Fig. 14.18 Fracture toughness of Sic whiskermullite composites with the crack propagating normal (L-type) and parallel (T-type) to the .predominant whisker direction&.
Table 14.10 Room temperature mechanical properties of Sic whiskers/Si,N, matrix composites45
Whisker content (Vol. Yo)
Fracture strength (MPa)
Fracture toughness (MPa mIn)
0 10 20 30
375 395 550 970
4.0 4.9 7.0 6.4
I
Sic Whisker Fraction ( ~ 0 1 % )
I
a
apparently above this fiber content.An alternative form of Sic whiskers was used to fabricate Si,N, matrix composites by Shalek et ~ 1 . 4 This ~. whisker is formed by the Vapor-Liquid-Solid p r o ~ e s s ~The ~ , effect ~ ~ . of whisker content and processing temperature on the elastic modulus, fracture strength and fracture toughness are shown in Figs. 14.19, 14.20 and 14.21 respectively. There is an apparent critical processing temperature above 1600°C to achieve maximum attainable strength and toughness properties in these composites. Continuous fiber ceramic matrix composites are more likely to obey the rule of mixtures prediction for elastic and strength properties
I I J I I I I I I I I I I IO
I
SiCw-mulIite
Sicw - mullite
2000
*
Properties
327
especially where minimum fiber damage and porosity occurs. The predictable increase in flexural strength for a Nicalon fiber/borosilicate glass composite is seen in Fig. 14.2249.
SIC Whisker -hobpressed Si3 N, matrix composite
SIC Whisker -hot-pressed Si3 N, matrix composite
13
300
: Hot-pressed at 1750°C B
: Hot-pressed
at 1850°C 1600°C
A : Hot-pressed at
-0
1 . . . . . . . . 1 5
10
IS
PO
25
30
35
40
Volume Yo S i c whiskers
Fig. 14.19 Elastic modulus of Sic whisker-Si,N4 composites processed at various pressing temperature~~~.
: Hot-pressed at 1750OC :Hot-pressed at 1850°C A : Hot-pressed at 1600°C 0
4
.
.
.
.
1
5
30
10
15
25
20
35
30
40
Volume % SIC whiskers Si, N, matrix composite 1750°C 1850°C . 1600°C
E3m-
*
03
"
50 -,p
:
t m.
8
A A
I
m
Rule of mixtures prediction for SiC/Borosilicate class
2500
B
4-4
0)
Fig. 14.21 Fracture tougness of Sic whisker-Si,N4 composites processed at various pressing temperatures&.
A
2 2000 E
# ,
,x
F
5 1000 1000-
'7 PI 0
5
10
IS
20
25
30
35
40
Volume Yo Sic whiskers
Fig. 14.20 Fracture strength of Sic whisker-Si,N4 composites processed at various pressing temperatures".
'9 ,
,4 /**4
500-
I
',
, ,
5 1500 1500-
A
,
, 0
/
/
/
O f '
.1
'
.
'
'
'
'
.2 .3 .4 .5 .6 .7 Fibre volume fraction
' .8
'
.9
Fig. 14.22 Rule of mixtures strength for Sic fiber (NicalonTM)-borosilicate glass49.
328
Ceramic composites Crater Wear
14.6 APPLICATIONS
14.6.1 HIGH TEMPERATURE STRUCTURES
The most advanced demonstration of ceramic matrix composites for high temperature structures and components has been for Sic CVI infiltrated carbon and Sic (NicalonTM) fiber. Extensive application of these composites have been made by Societe Europeenne de Propulsion A demonstration twostroke 100 cc engine consisting of Sic-Sic piston, cylinder and cylinder head has run for ten hours at full load and a speed of 1500 rpm without lubricant. A list of fabricated and tested high temperature ceramic composite structures is given in Table 14.11. These components operated successfully under actual- or simulated service conditions. 14.6.2 TRIBOLOGICAL COMPONENTS (CUTTING)
The application of ceramic composites to cutting tool inserts has made a significant impact on machining of difficult-to-cut metals. The typical wear pattern in a cutting tool insert is shown in Fig. 14.23.Uniform wear due to the rubbing action of the metal work piece produces the flank and nose wear. Deeper wear patterns at the depth-of-cut and trailing edge region of the insert are due to the sharp edges
Notch
Nose Wear
Fig. 14.23 Typical wear pattern in cutting tool inserts.
of the chip. The abrasion of the rake by the broad face of the chip produces a crater. Material removal from the rake face can be by dissolution, adhesion and chemical reaction. Low chemical reactivity, hot hardness and wear resistance allows ceramics to minimize all these forms of wear. Unreinforced ceramics, while capable of high cutting speeds, suffer from unpredictable life due to low impact load tolerance
Table 14.11 Typical high temperature applications of ceramic matrix composites
Application
Combustion chamber Turbine blade Turbine wheel Turbine wheel Leading edge Nozzle Radiant burner tube
Composition
Operating temperature ("C)
Environment
Test condition
Reference
Sic-Sic
1200
Oxidizing
14 h running
50
Sic-Sic Sic-Sic C-Sic C-Sic C-Sic
1200 1150 1150 1400 1550
Air-kerosene Air-kerosene Air-kerosene Oxidizing Oxidizing
50 50 50 51 52
Nextel 312-Sic
1150
Air-natural gas
Thermal cycle 400-1200°C 1 h up to 55 000 rpm 70 000 rpm Several hours Maximum thermal gradient 60"C/mm 18 month operation
52
Applications 329 or poor thermal shock properties. Reinforcing ceramics especially by whiskers is an effective technique to defeat these limitations. The best known example of this application is the Sic whisker reinforced A1,0, insert. The commercially available composition contains about 30% whiskers and is designated WG300 by Greenleaf Corp., Saegertown, PA, USA. The range of machining parameters for the Sic whisker alumina compared to other conventional and advances cutting tool materials is shown in Fig. 14.2454.Carbide cutting tools are limited to cutting speeds below 100 m/min while the ceramic cutting tools range from over 100m/min to 450m/min. The Sic whisker reinforced composite can be seen to provide the largest range of machining parameters for the machining of nickel based superalloys compared to other advanced cutting materials such as Sialon and Tic particulate reinforced alumina. The Sic whisker reinforced alumina (WG300) also has significantly greater tool life and allows much greater rates of metal removal for Inconel 718 compared to Tic particulate
reinforced alumina and cemented tungsten carbide as seen in Fig. 14.2555.
2
35 -
r 30c
S
c
'E 25> K .-
+ 200
K ._
3
15-
0
$ - 10m c L
2
50-
Fig. 14.25 Comparison of tool life and metal removal rates between various cutting tool materiais55.
-
0.25
Al,O,-SiC(w) Sialon
C a
0.20
FE
i
-
E 0.15
U
(I)
2 0.10
0.05 0
100
200
300
400
500
Cutting speed (m/min)
Fig. 14.24 Approximate range of machining parameters allowed by various cutting tool materials%.
330 Ceramic composites 15. Mecholsky, J.J., Engineering research needs of advanced ceramics and ceramic-matrix com1. Karasek, K.R., Bradley, S.A. Dormer, J.T., Yeh, H. posites, Am Ceram. SOC. Bull., 1989, 68(2), C., Schienle, J.L. and Fang, H.T., 367-375. Characterization of silicon carbide whiskers, I. 16. Maloney, L.D., Make way for 'Engineered Am. Ceram. SOC.,1989, 72, 1907-1913. Ceramics,' Design News, March 13, 1989, 45(5), 2. Homeny, J., Whisker reinforced ceramics, in: 64-71. Ceramic Matrix Composites (ed R. Warren), New 17. Naslain, R., CVI Composites, in: Ceramic Matrix York, Chapman and Hall: 1992, p.245. Composites (ed. R. Warren), New York: 3. Romine, J.C., New high temperature ceramic Chapman and Hall, 1992, p. 199. fiber, Ceram. Eng. Sci. Proc., 1987,8, 755. 18. Buckley, J.D., Carbon-carbon: an overview, Am. 4. Holtz, A.R. and Grether, M.F., High temperature Ceram. SOC.Bull., 1988, 67(2) 364-338. properties of three Nextel ceramic fibers, 19. Besmann, T.M., Sheldon, B.W. and Kaster, M.D., Presented at 32nd International SAMPE Temperature concentration dependence of Sic Symposium and Exhibition at Anaheim deposition on Nicalon fibers, Surf. Coatings Convention Center, April 6-9, 1987. Technol., 1990,43144, 167-175. 5. Bunsell, A.R., Development of fine ceramic 20. Stinton, D.P., Besmann, T.M. and Louden, R.A., fibers for high temperature composite, Materials Advanced ceramics by chemical vapor deposiForum, 1988,11, 78. tion techniques, Am. Ceram. SOC. Bull., 1988, 67 6. Ishikawa, T., Recent developments of the Sic (2), 350-355. fiber Nicalon and its composites, including 21. Warren, R. and Lundberg, R., Principles of properties of the Sic fiber Hi-Nicalon for ultrapreparation of ceramic composites, in: Ceramic high temperature, Compos. Sci. Tech., 1994, 51, Matrix Composites (ed R. Warren), New York: 135-144. Chapman and Hall, 1992, p35. 7. Okamura, K., Ceramic fibers from polymer pre- 22. van Konijnenburg, J.T., Siskens, C.A.M. and cursors, Composites, 1987,18(2), 107-127. Sinnema, S, Practical designing aspects of engi8. Flock, W.M., Characterization and processing neering ceramics, in: Designing with Structural interactions, in: Ceramic Processing Before Firing, Ceramics, (ed R.W. Davidge and H.M. Van de (ed. George Y. Onoda, Jr. and Larry L. Hench) Voorde), New York: Elsevier Science Publishing New York: John Wiley and Sons, 1978, p.31. Co, Inc., 1991, p.98. 9. Laynge, F.F., Lam, D.C.C., Sudre, O., Flinn, B.D., 23. Pigott, M.R., Load Bearing Fibre Composites, Folsom, C., Velamakanni, B.V., Zok, F.W. and Oxford: Pergamon Press, 1980, p.106. Evans, A.G., Powder processing of ceramic 24. Aveston, J., Cooper, G. and Kelly, A,, Single and matrix composites, Mater. Sci. Eng., 1991, A144, multiple fracture, in: The Properties of Composites, 143-152. Guildford: IPC Science and Technology Press, 10. Kragness, E.D., Processing and mechanical 1971, p.15. behavior of tape cast and laminated silicon car- 25. Halpin, J.C. and Kardos, J.L. The Halpin-Tsai bide whisker/alumina composites, M.Sc. equations: a review. Polym. Eng. Sci., 1976,16(5), Thesis, Pennsylvania State University, 1988. 344-352. 11. Prewo, K.M., Brennan, J.J. and Layden, G.K., 26. McCullough, R.L., Wu, C.T., Seferis, J.C. and Fiber reinforced glasses and glass ceramics for Lindenmeyer, P.H. Predictions of limiting high performance applications, Am. Ceram. SOC. mechanical performance for anisotropic crysBull. 1986,65(2), 305-322. talline polymers. Polym. Eng. Sci. 1976, 16(5), 12. Hillig, W.B., Melt infiltration approach to 371-387. ceramic matrix composites, 1. Am. Ceram. SOC., 27. Schapery, R.A. Thermal expansion coefficients 1988, 71(2), C9W99. of composite materials based on energy princi13. Newkirk, M.S., Lesher, H.O., White, D.R., ples. J. Compos. Mater. 1968, 2(3), 280404. Kennedy, C.R., Urquart, A.W. and Claar, T.D., 28. Kragness, E.D., Amateau, M.F. and Messing, Formation of LanxideTMceramic composite G.L., Processing and characterization of lamimaterials, Ceram. Eng. Sci. Proc., 1987, 8,879. nated Sic whisker reinforced A1203, J. Compos. 14. Newkirk, M. S., Urquart, A.W., Zwicker, H.R. Mater., 1991, 25(4), 416432. and Brevel, E., Formation of LANXIDETM 29. Paul, B., Predictions of elastic constants of mulceramic composite material, J. Mater. Res., 1986, tiphase materials, Trans. Met. SOC.,AIME, pp. 1(1),81-88.
REFERENCES
References 331 3 6 4 1 (February 1960). 30. Pister, K.S. and Dong, S.B., Elastic bending of layered plates, 1. Eng. Mech. Din, ASCE, 1-10 (October 1959). 31. Reissner, E. and Stavsky, Y., Bending and stretching of certain types of heterogeneous aelotropic plates, J. Appl. Mech., 402408 (September 1961). 32. Amateau, M.F., Properties of laminated ceramic composites, The 37th Sagamore Army Materials Research Conf Proc., (ed. D.J. Viechnicki), Watertown, Mass: Materials Technology Laboratory, 317-338, October, 1990. 33. Evans, A.G. and Davidge, R.W., A biaxial stress method for the determination of the strengh of sections cut from glass containers and the size of critical Griffith flaws, Glass Tech. 1971, 12(6), 148-154. 34. Evans, A.G. and McMeeking, R.M., On the toughening of ceramics by strong reinforcements, Acta Met., 1986, 34, 2435-2441. 35. Becher. P, Hsueh, C.-H., Angelini, P. and Tiegs, T.N., Toughening behavior in whisker-reinforced ceramic matrix composites, J. A m . Ceram. SOC.,1988,71,1050-1061. 36. Warren, R., Fundamental aspects of the properties of ceramic-matrix composites, in: Ceramic Matrix Composites (ed by R. Warren), New York: Chapman and Hall, 1992, p. 82. 37. Faber, K.T and Evans, A.G., Crack deflection process, Part I and II., Acta Met., 1983, 31, 565-584. 38. Hutchinson, J.W., Crack tip shielding by microcracking in brittle solids, Acta Met., 1987, 35, 1605-1619. 39. Vaughn, W.L., Homeny, J. and Ferber, M.K., Mechanical properties of silicon carbide whisker/alumina oxide matrix composites, Ceram. Eng. Sci. Proc., 1987, 8(7-8), 848-859. 40. Tiegs, T.N. and Becher, P.F., Sintering of Al,O,-SiC whisker composites, Bull. A m . Ceram. SOC.,1987,66,347-352. 41. Porter, J.R., Langej, F.F. and Chokshi, A.H., Processing and creep performance of Sic whisker reinforced A1,0,, Bull. Am. Ceram. SOC., 1987,66,343-346. 42. Lio, S., Watanabe, M., Matsubara, M. and Matsuo, Y., Mechanical properties of alumina/silicon carbide whisker composites, J. Am. Ceram. SOC.,1989,72,1880-1884. 43. Becher, P.F., Tiegs, T.N., Ogle, J.C. and Warwick, W. H., Toughening of ceramics by whisker reinforcement, in: Fracture Mechanics of Ceramics,
VoI. 7, Composites, Impact, Statistics and HighTemperature Phenomenon, (ed. R.C. Bradt, D.P.H. Hasselman, A.G. Evans and F.F. Lange), New York Plenum Press, 1986, pp. 61-72. 44. Wu, M., Messing, G.L. and Amateau, M.F., Laminate processing and properties of oriented Sic whisker-reinforced composites, Ceramic Transactions, Vol. 19, 1991, Westerville, OH: American Ceramic Society, 665476,1991. 45. Bulgan, S.T., Baldoni, J.G. and Huckabee, M.L., Si,N4-SiC whisker composites, Bull. Am. Ceram. SOC.,1987,66,347-352. 46. Shalek, P.D., Petrovic, J. J., Hurley, G.F. and Gac, ED., Hot-Pressed Sic Whisker/Si,N, Matrix Composites, A m . Ceram. SOC., 1986, 65(2), 351-356. 47. Milewski, J.V., Gac, ED., Petrovic, J.J. and Skaggs, S.R., Growth of beta-silicon carbide whiskers by the VLS process, J. Mater. Sci., 1985, 20,1160-1166. 48. Petrovic, J.J., Milewski, J.V., Rhor, D.L. and Gac, ED., Tensile mechanical properties of Sic whiskers, 1.Mater. Sci., 1985,20, 1167-1177. 49. Dawson, D.M., Preston, R.F. and Purser, A., Fabrication and materials evaluation of high performance aligned ceramic fiber-reinforced, glass-matrix composite, Ceram. Eng. SOC.Proc., 1987,8,815-821. 50. Heraud, L.P. Spriet, 'High toughness C-SIC and Sic-Sic composites in heat engines', in Whiskerand Fiber-Toughened Ceramics, Proceedings of an International Conference, (ed. R.A. Bradley, D.E. Clark, D.C. Larsen and J.O. Stiegler), International Metals Park, OH: ASM, 1988. 51. Melchior, A.B., Pouliquen, M.F., Soler, E., Thermostructural composite materials for liquid propellant rocket engines, Paper AIAA-87-2119, AIAA/SAE/SME/ASEE, 23rd Joint Propulsion Conference, June 29-July 2, 1987, San Diego, CA, American Institute of Aeronautics and Astronautics, Washington, DC. 52. Suacereau, D. and Beaurain, A., Demonstration of carbon-silicon carbide Novoltex reinforced composite nozzle on a LH2-LOx, Engine, Paper AIAA-90-2180, AIAA/SAE/SME/ASEE, 26rd Joint Propulsion Conference, July 16-18, 1990, Orlando FL, American Institute of Aeronautics and Astronautics, Washington, DC. 53. Richards, R.E., Bodkins, D. W. and Copes, S. J., Progress toward a cost effective thin wall RBT, in Energy Technology: Processings of The Energy Technology Conference, Vol. 15, Energy Technology Conference, February 17-19, 1988,
332
Ceramic composites
Washington, DC, Government Institute Inc. Washington DC, 749-756,1988. 54. Billman, E.R., Mehrota, P.K., Shuster, A.F. and Beegley, C.W., Machining with A1203-Sic whisker cutting tools, Bull. Am. Ceram. Soc, 1988,67,1016-1019.
55. Rhodes, J.F., Whisker reinforced ceramic composites, in Proc. Fifth Ann. Conf. Maferials Technology, Materials Technology Center, Southern Illinois University, Carbondale, IL, 205-219,1988.
CARBON-CARBON COMPOSITES
15
John D. Buckley
15.1 INTRODUCTION
n
Carbon-carbon (CC) materials are a generic class of composites similar to the graphite/epoxy family of polymer matrix composites. These materials can be made in a 1-D 2-D wide variety of forms, from one-dimensional to n-dimensional, using unidirectional tows, tapes, or woven cloth (Fig. 15.1). Because of their multiformity, their mechanical properties can be readily tailored (Table 15.1). Carbon materials have high strength and stiffness potential as well as high thermal and chemical stability in inert environments. These materi3-D n-D als must, however, be protected with coatings and/or surface when used in an OXi- Fig. 15.1 Multiformity and general properties of dizing environment. carbon-fiber and carbon-matrix composites.
Table 15.1 General properties of carbon-carbon composites Ultimate tensile strength Modulus of elasticity Melting point Thermal conductivity Linear thermal expansion Density
>276 MPa
>40 000 psi
>69 GPa
>lo7 psi
>41OO0C ~11.W 5 m-' K-'
7412°F 6.64 h ft "F
~ 1 .x110"OC
6.1 x W7"F
<2990 kg m-3
186.6 lb/ft3
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
The development of CC materials began in 1958 and was nurtured under the US Air Force space plane program, Dyna-Soar and NASA's Apollo projects. It was not until the Space Shuttle Program that CC material systems were intensively researched. The criteria that led to the selection of CC composites as a thermal protection system were based on the following requirements: 1. maintenance of reproducible strength levels at 1650°C (3002°F); 2. sufficient stiffness to resist flight loads and large thermal gradients; 3. low coefficient of thermal expansion to minimize induced thermal stresses; 4. oxidation resistance sufficient to limit strength reduction;
334 Carbon-carbon composites 5. tolerance to impact damage; 6. manufacturing processes within the state of the art. Carbon-carbon composites consist of a fibrous carbon substrate in a carbonaceous matrix. Although both constituents are the same element, this fact does not simplify composite behavior because the state of each constituent may range from carbon to graphite. Crystallographic carbon, namely graphite, consists of tightly bonded, hexagonal arranged carbon layers that are held together by weak van der Waals forces. The single crystal graphite structure is illustrated in Fig. 15.2 (Bokros, 1969). The atoms within the layer
plane or basal plane (a-b direction) have a covalent bond strength of =524 kJ/mol (Kanter, 1957), while the bonding energy between basal planes (c direction) is =7 kJ/mol (Dienes, 1952). The result is a crystal that is remarkable in its anisotropy, being almost isotropic within the basal plane but with c direction properties that differ by orders of magnitude. On a larger scale, carbon, in addition to its two well-defined allotropic forms (diamond and graphite), can take any number of quasicrystalline forms ranging continuously from turbostratic (amorphous, glassy carbon) to a highly crystalline graphite (Fig. 15.3) (Bokros, 1969).
C
E
a
Reference directions
Fig. 15.2 Tightly bonded, hexagonally arranged carbon layers held together by weak van der Waals forces.
I
:rdo02
LC
Fig. 15.3 Comparison of (a) carbon turbostratic structure with (b) 3-D graphite lattice (Bokros, 1969).
Carbon fibers 335 The anisotropy of the graphite single crystal encompasses many structural forms of carbon. It ranges in the degree of preferred orientation of the crystallites and influences porosity, among other variables. A broad range of properties is the result of this anisotropy. In CC composites, this range of properties can extend to both constituents. Coupled with a variety of processing techniques that can be used in the fabrication of CC composites, great flexibility exists in the design of and the resultant properties obtained from CC composites. The wide range of properties of carbon materials can be shown when comparing the tensile modulus of commercially manufactured carbon fibers that range from 27.6 GPa (4 x 106psi) to 690 GPa (100 x lo6psi). In fabrication, the fibers can be used in either continuous or discontinuous form. The directionality of the filaments can be varied ranging from unidirectional lay-ups to multidirectional weaves. Fiber volume fraction constitutes another variable. The higher the volume fraction of a specific high-strength fiber in a matrix, the greater the strength of the composite. The matrix can be formed via two basic approaches: (1) through the carbonization of an organic solid or liquid, such as a resin or pitch, or (2) through the chemical vapor deposition (CVD) of carbon from a hydrocarbon. A range of carbon structures can be obtained by either approach. Finally, heat treatment of the composite material at graphitization temperatures offers additional variability to the properties that can be obtained. Typically, there is an optimum graphitization temperature at which the highest strength can be obtained for a given composite composition of fiber and matrix (Edie et al., 1986; Stoller et al., 1974). 15.2 CARBON FIBERS
The properties of carbon fibers can vary over a wide range depending on the organic precursor and processing conditions used. At present, graphite fibers are produced from
three precursor materials: rayon, polyacrylonitrile (PAN)and petroleum pitch. Fibers having a low modulus (27.6 GPa (4 x 106 psi)) are formed using a rayon precursor material that may be chemically pretreated by a sequence of heating steps. First, the fiber is heated to >400"C (752°F) to allow cellulose to pyrolyze (decomposition or chemical change during thermal conversion of organic materials to carbon and graphite). Carbonization (continued heating of organic material to >lOOO"C (1832°F) to initiate ordering of the carbon structures produced by pyrolysis) is completed more rapidly at >lOOO"C (1832°F).Upon completion of carbonization, the fiber is graphitized (continued heating of carbonized organic materials to the 2000-3000°C (3632-5432°F) range of produce 100% graphite-ordered crystal structure) by heating to >2000"C (3632°F);the fiber is now, for all practical purposes, 100% carbon. High-modulus carbon fibers from rayon precursors are obtained by the additional process of stretching the carbon fibers at the final heat treatment temperature. Nigh-modulus (344 GPa (50 x 106 psi)), high-strength (2.07 GPa (300 x lo3 psi)) carbon fibers are typically made from PAN or, in some cases, mesophase pitch precursors. These fibers are processed similarly in a three-stage operation (Fig. 15.4) (Diefendorf, 1987).The PAN fibers are initially stretched from 500-1300% and then stabilized (cross-linked) in an oxygen atmosphere at 200°C (392°F)to 280°C (536°F) under tension. Carbonization of the fibers is conducted between 1000°C (1832°F)and 1600°C (2912°F). Finally, graphitization is accomplished at >2500"C (4532°F). Mesophase pitch fibers undergo the same processing procedure as PAN fibers but do not require an expensive stretching process during heat treatment to maintain preferred alignment of crystallites (Fig. 15.4) (Diefendorf, 1987). Control of fiber shape has resulted in improved fiber strength (4.1 GPa, 600000 psi) (Cogburn et al., 1987) when produced from melt-spun, mesophase petroleum pitch (Fig. 15.5) (Cogburn et al., 1987). Round fibers using the same method
336 Carbon-carbon composites
PAN process
1
Pitch process
tCarbonize z z I I Graphitize zim Spool
Epoxy sizing
Surface treatment
Fig. 15.4 Carbon fiber production using PAN and pitch processes (Diefendorfer, 1987).
Hydraulic piston Cartridge housing Heating collar Melted-pitch precursor Spinnerette
Melt-pressure indicator
Melt-spun carbon filaments
Wind-up bobbin
Fig. 15.5 Melt spinning apparatus used to produce noncircular carbon fibers (Cogbum et al., 1987).
Carbonfibers 337 had a strength of 2.1 GPa (300 x lo3 psi) (Edie et al., 1986).Of the shapes studied, the C-shape and hollow fibers were found to be superior in strength to round solid and trilobal cross sections (Edie et al., 1986; Cogburn et al., 1987). 15.2.1 CARBON FIBERS IN CARBON MATRIX
2. achieve a more nearly isotropic material; 3. increase the composite interlaminar tensile strength; 4. along with continuous filament substrates, obtain a stronger composite by providing additional nucleation sites that serve to reduce composite porosity.
The most widely used starting materials are a carbonized, rayon felt substrate with a pyrolytic carbon matrix and short, chopped fibers in a pitch-based matrix. Felt is produced through the mechanical carding of viscous rayon fibers to produce a continuous web of fibers. The webs are folded one on top of another to produce a batt. The batts are then cut, stacked and needled to produce the 15.2.2 DISCONTINUOUS FIBER COMPOSITES required felt. The rayon felt is subjected to a controlled carbonization cycle in an inert Fabrication of discontinuous fiber composites atmosphere or vacuum; the maximum temperuses short carbon fibers combined with either ature determines such factors as shrinkage, a pyrolytic carbon or pyrolyzed organic weight loss and chemical composition of the matrix. This approach to CC composites genfelt. A maximum carbonization temperature of erally does not have true fiber reinforcement 1200°C (2192°F) is a nominal standard; the as an objective. Rather, discontinuous fiber length of the carbonization cycle and rate of substrates have been used to: temperature rise are dictated by the thickness 1. increase fabrication capability of large-scale of the felt. Carbon content in the fibers is ~ 9 8 % .Carbon-arbon composites have also structures; Addition of a matrix to carbon fiber, either through the carbonization of an organic precursor or by the deposition of pyrolytic carbon, is conducted at 800°C (1472°F) to 1500°C (2732°F).Subsequent heat treatment of the composite material may involve temperatures to 3000°C (5432°F).
Fig. 15.6 Models of fiber arrangements for four short-fiber fabrication techniques: (a) flocking lay-up, (b) pulp molding, (c) isotropic casting, and (d) spray lay-up (Cook, Lambdin and Trent, 1970; Lambdin, Cook and Marrow, 1969; Lambdin and Cook, 1971).
338 Carbon-carbon composites
been fabricated from short carbon fibers using isotropic casting, flocking lay-up, spray lay-up and pulp-molding techniques (Fig. 15.6) (Cook, Lambdin and Trent, 1970; Lambdin, Cook and Marrow, 1969; Lambdin and Cook, 1971.).The rationale for using these short fibers is to reduce composite anisotropy (Lambdin, Cook and Marrow, 1969). 15.3 CONTINUOUS FIBER COMPOSITES
Continuous filament substrates reflect the properties of high-strength filaments or achieve a high degree of preferred orientation on the macroscale of the matrix. The fabrication complexity for continuous-filament substrates is determined by two parameters: (1)the directionality of the filaments, and (2) the amount of layer interlocking achieved in the substrate. Filament winding of unidirectional tapes can be used to achieve a highly
oriented substrate, usually with no interlocking between layers. Woven fabrics are used to form a two-dimensional laminate with no interlocking between layers. Helical filament winding, which is directional, results in continuous, adjacent layer interlocking. Multilayer locking is achieved through complex weaving patterns or yarn placement resulting in 'multidirectional' substrates (Fig. 15.7). 15.4 CHEMICAL VAPOR DEPOSITION
The CVD of carbon from a hydrocarbon gas within a substrate is a complex process. Various techniques have been applied to infiltrate various fiber substrates including isothermal thermal gradient (Pierson, 1968), pressure gradient (Kotlensky and Pappis, 1969) and pressure pulsation (Beatty and Kipplinger, 1970).The first two have been the
Fig. 15.7 Interlocking approaches of continuous filament substrates: (a) tape wrapped, shingle; (b) filament wound, helix; and (c) multidimensional.
Carbonized organic composites 339 most extensively used. The isothermal technique is illustrated in Fig. 15.8. The substrate is radiantly heated by an inductively heated susceptor so that the gas and substrate are maintained at a uniform temperature. Infiltration is normally accomplished at 1100°C (2012°F) and at reduced pressures (6 kPa (50 torr)) with the flow rates primarily determined by the substrate surface area. This technique produces a crust on the outer surfaces of the substrate, thus requiring machining and multiple infiltration cycles. In the thermal gradient technique (Fig. 15.9), the part to be infiltrated is supported by a mandrel that is inductively heated. Therefore, the hottest portion of the substrate is the inside surface, which is in direct contact with the mandrel. The outer surface of the low-density substrate is exposed to a cooler environment and results in a temperature gradient through the substrate thickness. Surface crusting is eliminated because the deposition rate is greater on the heated fibers near the mandrel, whereas the cooler outer fibers receive little or no deposit. Under proper infiltration conditions, the carbon is first deposited on the inside surface and, in a continuous process, progresses radially through the substrate as the densified substrate itself becomes inductively heated. Infiltration is normally accomplished at atmospheric pressure with a mandrel heated to approximately 1100°C (2012°F) (Theis et al., 1970).
Hydrocarbon gas
2(-Carrier gas
Original fiber substrate
Fig. 15.8 Isothermal chemical vapor deposition to infiltrate fibrous carbon substrate.
15.5 CARBONIZED ORGANIC COMPOSITES
Carbonized organic composites have fabrication procedures that are similar to those of conventional fiber-reinforced, resin-laminating techniques. The starting material is usually a prepregged fabric or yarn (a fabric impregnated with a matrix material in a tacky state). These precursor materials are staged nominally at approximately 100°C (212°F) to achieve the desired degree of tack and flow of the resin. A laminate is then constructed and cured under pressure. Curing temperatures
Fig. 15.9 Thermal gradient chemical vapor deposition.
340 Carbon-carbon composites
AprepreK+ Cut, lay-up,
3 t o 5 times
I
cure
I
Fig. 15.10 Fabrication steps involved in manufacture of 2-D carboniarbon part impregnated with tetraethylorthosilicate (TEOS).
normally range from 125°C (257°F) to 175°C (347°F) with curing pressures on the order of 2.76 MPa (400 psi). The reinforced resin laminate is then post-cured at 200°C (392°F) to 275°C (527°F).As pyrolysis is initiated, shrinking occurs as the organic phase decomposes. Simultaneously the release of vapors from pyrolysis expands the composite material. A slow release of these volatile by-products is required to minimize structural damage to the char. Finally, as higher temperatures are reached, thermal expansion of the carbon char itself occurs after pyrolysis is complete. After the initial carbonization, the material is then subjected to a series of reimpregnation and carbonization cycles until the desired density or the maximum density is achieved. The reimpregnation process is usually conducted under vacuum and pressure to aid in maximizing the pore filling. If graphitization is desired, the high-temperature heat treatment may be used after each carbonization step or at the end of the reimpregnation and recarbonization cycles.
To summarize, a typical manufacturing cycle of a 2-D CC part is shown in Fig. 15.10.First, a woven graphite fabric that is preimpregnated with phenolic resin is laid up as a phenolicgraphite laminate in a mold and is autoclave cured. Once cured, the part is pyrolyzed to form a carbon matrix surrounding the graphte fibers. The part is then densified by multiple furfural alcohol reimpregnations and pyrolyzations. The resulting CC part then is ready for use in inert or vacuum environments. This process is very time consuming. A single pyrolysis may take >70 h in a low-temperature, inert-atmosphere furnace. Although CC materials can withstand temperatures >3000"C (5432°F) in a vacuum or in an inert atmosphere, they oxidize and sublime when in an oxygen atmosphere at 600°C (1112°F). To allow for use of CC parts in an oxidizing atmosphere, they must be compounded with materials that produce oxidation-protective coatings through thermochemical reaction with oxygen at >2000"C (3632°F) (Buckley, 1967) or they must be coated
Manufacturing 341 and sealed to protect them (Strife and Sheehan, 1988). For applications such as the Space Shuttle CC leading edges and nose caps, surfaces are converted to silicon carbide in a high-temperature diffusion-coating process (Fig. 15.10). Because of differences in thermal expansion between the silicon carbide and the CC part, the coating develops microcracks when the part is cooled from the coating temperature. To maintain oxidation protection on space vehicles such as the Space Shuttle, cracks are impregnated with tetraethylorthosilicate (TEOS).The TEOS process leaves silica in all of the microcracks, greatly enhancing the oxidation protection of the CC substrate. Current improvements being developed for oxidation protection of the CC Space Shuttle components are additions of low-temperature glass formers that enhance the sealing capability of the existing coating-TEOS system. 15.6 MANUFACTURING
The fabrication process of the Space Shuttle Orbiter nose cap and wing leading edge components (Fig. 15.11) (Curry, Scott and Webster, 1979) is a multi-step process typical of the technology used to produce CC composites. The process steps are illustrated in Figs. 15.12-15.16. Initial material lay-up is similar to conventional practices with fiberglass-reinforced plastic parts. Square-weave graphite fabric impregnated with phenolic resin is laid-up in an epoxy/fiberglass mold cavity shaped to the desired configuration (Fig. 15.12) (Curry, Scott and Webster, 1979).Lay-up thickness for these components varies from 19 plies in the external skin and web areas to 38 plies in the attachment locations. Upon completion of layup, the part is vacuum-bagged and cured in an autoclave to 150°C (300°F) for 8 h (Fig. 15.12). The cured part is rough trimmed, X-rayed and ultrasonically inspected for irregularities following the cure cycle. Post-cure of the component involves placing the part in a graphite restraint fixture loading it into the
RCC Seal strips (22) LH (22) RH
Wing L.E. RCC panels
Nose cap (1) RCC Seal strip (1) LH, (1) RH and (3) Lower RCC ExDansion seal (1) LH, (1) RH and (3) Lower
Fig. 15.11 Leading edge structural subsystem (Curry, Scott and Webster, 1979).
furnace and submitting it to a 7-day cycle during which it is taken to 260°C (500°F) very slowly to avoid distortion and delamination (Fig. 15.12). The next step is initial pyrolysis as shown in Fig. 15.13 (Curry, Scott and Webster, 1979). Pyrolysis tooling composed of graphite restraining fixtures containing the part are loaded into a steel retort that is packed with calcined coke. The retort and its contents then undergo a 70 h pyrolysis cycle at 815°C (1500°F)converting the phenolic resin to a carbon state. During pyrolyzation, the resin forms a network of interconnected porosity for the escape of volatile matter. This stage is extremely critical since, during controlled charring of the cured resin matrix, the parts are weak and delamination can easily occur if adequate escape paths and time are not ensured. After this initial pyrolysis cycle, the carbon is
342 Carbon-carbon composites 2250 1650
w 1100 0
I
j
,
I
2
540 485
2
g% :
u1 320
260 200 150'
UJ
8
; I I
; n
MATERIAL COLD I STORAGE I
j
I
I
CUT & LAY.UP CLOTH PLIES
0 0 0
BAG (MYLAR) APPLYVACUUM CHECK FOR LEAKS
AUTOCLAVE CURE
0
I
0 0
0
&
REMOVE BAG ROUGH T R I M DRILLHOLES X.RAY AND/OR ULTRASONIC
LOAD IN RESTRAINING FIXTURE
z
r , h
I\
I:
POST CURE
DEBULK
Fig. 15.12 Lay-up and cure cycle (Curry, Scott and Webster, 1979). -
2205
2250
1650
g iioc F
540
485 5 430 0 370
320 ~
260
a zw
x
C3 150 95 40 -2c
LOAD IN
LOADIN PYROLYSIS RETORT PACK WITH GRAPHITE CALCINED COKE
0
CLEAN INSPECT
I
Fig. 15.13 Initial pyrolysis (Curry, Scott and Webster, 1979).
AUTOCLAVE CURE
msi CURE j
VACUUM CHAMBER IMPREG WITH FURFURYL ALCOHOL IMPREGNATION CYCLE ITMREE 131 TIMESI
Fig. 15.14 Densification impregnation and cure (Curry, Scott and Webster, 1979).
Manufacturing 343 designated reinforced carbon-carbon-0 (RCC-0), a state in which the material is extremely light and porous with a flexure strength of 21-24 MPa (3000-3500 psi). 15.6.1 DENSIFICATION
2250
w
P
J I
F
5 w
a
Densification for these shuttle parts is accomplished in three impregnation and pyrolysis cycles (Fig. 15.14) (Curry, Scott and Webster, 1979).Each part is loaded in a vacuum chamber impregnated with furfural alcohol followed by a 2-hour cure period in the autoclave at approximately 150°C (300"F), followed by a post-cure for 32 h to 200°C (400°F). This cycle is followed by a 70-hour 815°C (1500°F) pyrolyzation that is shown in Fig. 15.13.After three impregnation/pyrolyzation cycles, the material is designated RCC-3 with an increased flexure strength of =124 MPa (18000 psi) at room temperature.
I
1650 ,
s
540 485 430 370 320 260 200 150 95 40 -20
CLEAN
'
COATING PREPARATION -X.RRY INRETORT SURFACE COATING ! -ULTRASONIC :e FINALTRIMIDRILL .1 DIM. INWECT
: ,
CLEAN INSPECT -X.RAY - ULTRASONIC - DIMENSIONAL
1 I
Fig. 15.15 Coating cycle (Curry, Scott and Webster, 1979).
15.6.2 COATING
To allow for use of CC composites at elevated temperatures above 2000°C (3632°F) in an oxidizing atmosphere, it is necessary to apply protective coatings to structural components. The oxidation inhibition process consists of two steps: (1)diffusion coating the CC component and (2) applying a sealer to the surface. The coating process (Fig. 15.15) (Curry, Scott and Webster, 1979) used in protecting the CC shuttle components starts with the blending of the constituent powders: 10% alumina, 30% silicon and 60% silicon carbide. This mix is packed around the CC structural component in a graphite retort and loaded into a vacuum furnace where it undergoes a 16-hour cycle that includes drying at 315°C (600°F) and the coating reaction to 1650°C (=3000"F) in an argon atmosphere. The powder characteristics, constituents, formulations and the manner in which the powders are packed around the part are important factors that govern the chemical reactions at the high processing temperatures, the degree of consol-
idation and sintering of the powders. During the process, the outer layers of the CC substrate are converted to silicon carbide. The silicon carbide-coated CC composite part is removed from the retort, cleaned and inspected. During cool down from 1650°C (3000"F), the silicon carbide coating contracts slightly more than the carbon substrate, causing crazing (coating fissures). This crazing together with the inherent material porosity provides paths for oxygen to reach the carbon substrate. To obtain increased useful life of this CC structural component, it is necessary to add an additional oxidation inhibitor. The final process used to provide oxidation protection to this type of CC structure involves impregnating (Fig. 15.16) (Curry, Scott and Webster, 1979)this component with TEOS. The part is covered with a mesh, placed in a vacuum bag and the bag is filled with liquid TEOS. A 5-cycle TEOS impregnation is then performed with the bagged part. After the fifth TEOS cycle, the part is removed from the bag
344 Carbon-carbon composites W
9 9 5 8
320 260 200
w
95
W
U
8
150 40
-20 45 TO
45 TO
6
o
MIN
w MIN
45 TO
m MIN
45 TO
2
.5
TO
HRS
a
2%
6
MIN
IHRS
HRS
. COATED PART PERSPEC 206.741
WEIGHT/ RECORD VERIFY -SRi PRi
-
COVER
wlm
MESH
INSTALL VACUUM FITTINGS -FILL TUEES COVER WITH MESH VACUUM BAG
-
CONNECTVACUUMAND F I L L LINES MIXTEOS ILIWIDI A N D FILL RESERVOIR
F I V E CYCLES EVACUATE AND BACKFILL REMOVE WITHTEOS BAG MIX CLEAN PART OVEN CURE 225" + 5O'F I 4 5 6 0 MINI FEED TEOS M I X A5 R E Q D TO COVER 5 T H CYCLE CURE U TO 2x H R MlNl COOL T O 1WoF
0
CURE % HR @ W ' F 6HR B W F
NDE X.RAY EDDY CURRENT -ULTRASONIC
-
-
-
0
WEIGH/ RECORD
Fig. 15.16 TEOS impregnation (Curry, Scott and Webster, 1979).
labeled Space Shuttle material, is the strength level of the reusable carbon-carbon (RCC) material used in the Space Shuttle thermal protection system. Even though this material is made with low-strength carbon fibers, its strength efficiency is superior to both superalloys -and ceramics at >lOOO°C (1832°F). 15.7 MECHANICAL PROPERTIES Development of advanced carbon-carbon The extreme thermomechanical requirements (ACC) composites has produced a material of the Space Shuttle have been the impetus for that is twice as strong as the CC composite evaluating properties of low-density CC. The first put on the Space Shuttle. The ACC mateuse of CC on the nose cap and leading edges of rial is made using woven carbon cloth. When the Space Shuttle makes it imperative to know unidirectional carbon fiber tapes are interplied as much as possible about all the characteris- with woven cloth to create a hybrid ACC, tics of this material. The effect of temperature strength in at least one direction can be on the ratio of tensile strength to density for increased by >345 MPa (>50 000 psi). Current several classes of high-temperature materials data on thermomechanical and thermochemiis shown in Fig. 15.17. The major advantage of cal properties of some of the advanced CC CC materials for high-temperature applica- systems show that material composition, oxitions is that they do not lose strength as the dation resistance, processing, joining and fiber use temperature is increased. This property is architecture are producing noticeable in contrast to other materials such as superal- improvements in CC materials and structures loys and ceramics. Figure 15.17 shows three (Curry, Scott and Webster, 1979; Buch, 1984; levels of CC strength efficiency. The first, Rummler and Sawyer, 1984; Ransone and and oven cured at 315°C (600"F), liberating all of the hydrocarbons. This procedure leaves silica (SiO,) in all of the microcracks and fissures greatly enhancing the oxidation protection of the CC structure.
Thermal properties 345 Temperature, OC 0
-
,550 I
1100 I
1650 I
High-strengthcarbon-carbon
800
1600 2400 Temperature, OF
3200
**%O
<-
160
4000
Fig. 15.17 Strength-to-density ratio for several classes of high-temperature materials.
Ohlhorst, 1984; Webb, 1985; Gray and Engle, 1985; Johnson and Finley, 1985; Sawyer and Moses, 1985; Maahs and Ransone, 1985; Ohlhorst and Ransone, 1985). CC components on the Space Shuttle are required to have adequate strength at design temperatures to withstand the aerodynamic loads of flight and to continue to do so for the operational life of the component. Minimum mechanical properties are guaranteed through statistical analysis of a data sampling having at least 99% probability and 95% confidence. The primary variables affecting the structural design allowables are temperature, material thickness, coating thickness, biaxial stress conditions and substrate mass loss due to oxidation through the mission life of the component (Table 15.2) (Curry, Scott and Webster, 1979). Figure 15.18 (Curry, Scott and Webster, 1979) illustrates the typical effect of ply thickness on the allowable stress values for tension, bending, compression and shear used for design. As fabricated, room temperature mod-
ulus values for the TEOS material are shown in Fig. 15.19 (Curry, Scott and Webster, 1979). The effect of temperature on the as-fabricated tension strength properties is shown in Fig. 15.20 (Curry, Scott and Webster, 1979). As shown in Figs 15.17 and 15.20, the strength of CC composite material does not decrease significantly with temperature. Typically, above 1425°C (2600°F) there is an increase in strength. The effect of substrate mass loss through oxidation on tensile strength is shown in Fig. 15.21 (Curry, Scott and Webster, 1979). Mass loss results in a significant reduction in design allowable stress, emphasizing the value of the additional oxidation protection provided by the TEOS treatment. 15.8 THERMAL PROPERTIES
15.8.1 THERMAL OXIDATION
A critical requirement when using CC composites is the ability to withstand numerous
346 Carbon-carbon composites Table 15.2
Non-TEOS
Mechanical properties test
TEOS
-.
__
as fabricated
conditioned
as fabricated
conditioned
Flexure Tension Compression Shear Comer flexure Interlaminar tension Interlaminar shear Coefficient of thermal expansion Impact, etc. Simultaneous cycling Mission cycling Tension Flexure Compression
201 196 170 192 20 20 30 10 20
221 80 57 79
40 52 46 17 3 5
24 34 28 17 3 5
-
Total
-
50 19 6 10
-
-
9
6
-
-
-
3
18
3 3 3
36 24 27
-
-
868
609
175
135
-
COMPRESSION
NOTE BREAK IN SCALE
TENSION 7
IN-PLANE SHEAR
14
I 15
I
1
20
25
I 30
1 35
I 40
NUMBER OF PLIES
Fig. 15.18 Design allowables at room temperature as-fabricated (Curry, Scott and Webster, 1979).
Thermal properties 347
35
E
14
Y
7.0
3.5
r
T V P I C A L SECANT HOOVLUS
t 1
Gw
1
I 10
1 I5
1 25
1 4a
I
1 30
35
NUMBER OF PLIES
Fig. 15.19 Design allowables at room temperature as-fabricated (Curry, Scott and Webster, 1979).
thermal and thermomechanical loads during re-entries of the Space Shuttle into the earth's atmosphere. Although CC Space Shuttle components have an oxidation-inhibiting silicon carbide coating, they can lose mass over an extended temperature range without apparent surface recession. Photomicrographs of CC specimen surfaces show minute fissures and thermal microcracks, some of which terminate at the coating substrate interface. Specimens exposed to convective and radiant heat transfer tests micrographically have shown the presence of voids at the coating substrate surface. Tests to characterize the effects of these
___-
z0 . Y 2
E
35
2
0
flaws were performed over a wide range of pressures and temperatures in both plasma arc jets and radiant-heating test facilities. Arc jet tests on CC specimens ranged in temperature from 815°C (1500°F) to 1870°C (3400°F) and atmospheric pressures from 0.01 Pa to 0.10 Pa. Radiant-test conditions ranged from 420°C (800°F)to 1425°C (2600°F)and pressures ranging from 0.01 Pa to 1.0 Pa. Mass-loss data for the CC shuttle specimens exposed to the arc jet and radiant-heating tests are presented in Figs 15.22 and 15.23. Figure 15.22 (Curry, Scott and Webster, 1979) shows mass loss at 980°C (1800°F)and 0.05 Pa as a function of exposure
28 PLY
---_______________--------111 P L V
. -
v)
2 Iw
28
-
m
ss 2
< 2 1 -
I
I
ur .
I
I
I
I
Fig. 15.20 Design allowables as-fabricated (Curry, Scott and Webster, 1979).
I
1
348 Carbon-carbon composites
38 PLY 28 PLY 25 P L Y 1s P L Y
3
c
UJ
1.4
0
I
1
I
1
4.9
9.8
14.7
1U.6
MASS
LOSS
-
J 24.5
kdm’x 1P
Fig. 15.21 Design allowables conditioned (Curry, Scott and Webster, 1979).
P
n
39.0
NONTEOS \
/ /
n
z
-
19.6 -
P
-i
/ /
.x W
I 14.7
-
/
v) v)
s $ 4
N O N TEOS
9.8
-
-
9
.
8
W I T1 H TEOS
4.9
-
, , d
1
/
,’
8 I
a
2
I
I
4
6
EXPOSURE TIME
I
1
8
10
HOURS
Fig. 15.22 Mass loss comparison plasma arc jet environment (Curry, Scott and Webster, 1979).
0
’
/
I
/
P,
8/
&--e--B \WITH
2
9
d/
4
EXPOSURE T I M E
TEOS
8
6
10
HOURS
Fig. 15.23 Mass loss comparison radiant environment (Curry, Scott and Webster, 1979).
Applications 349 14.5 7
11.1-
,-
\
SPECIFIC HEAT
THERMAL CONDUCTIVITY PARALLEL TO PLY
'F ai.
B
E
1 8 i'
a
-THERMAL
5'8'
CONDUCTIVITY PERPENDICULAR TO PLY
2.9.
+T -OTAL
f
EMISSIVITY
-0 -28(1
-18
280
538
818
1094
1372
16sO
TEMPERATURE "C
Fig. 15.24 Reinforced carbon-carbon thermal properties (Curry, Scott and Webster, 1979).
thermal conductivity is dependent upon the mass loss experienced by the CC composite, resulting from subsurface oxidation. Results of thermal conductivity studies for shuttle CC composite shuttle materials are shown in Fig. 15.24 (Curry, Scott and Webster, 1979).To simplify thermal modeling, no differentiation has been made for conductivity variation resulting from the number of plies in the substrate. Results for conditioned specimens having a mass loss of 0-5 Pa (0.1 lb/ft2) suggest that thermal conductivity decreases with mass loss. Figure 15.24 also shows that neither specific heat nor emittance was affected by material or mass loss conditioning. 15.9 APPLICATIONS
An example of the state of the art in CC composite applications is a one-piece, bladed turbine rotor that, in service, is coated to prevent oxidation. The rotor offers higher temperature performance without cooling; low weight and use of low-cost, non-strategic materials (Miller and Grimes, 1982).Other gas turbine engine applications using CC composites include exhaust nozzle flaps and seals, augmenters, combustors and acoustic panels.
CC material systems using coatings, TEOS and additions to the basic CC recipe have improved the oxidation resistance of products made of CC composites by an order of magnitude. These composites are being used in products such as the nozzle in the F-100 jet engine afterburner, turbine wheels operating at >40 000 rpm, nonwetting crucibles for molten metals, nose caps and leading edges for missiles and for the Space Shuttle, windtunnel models and racing car and commercial disk brakes (Klein, 1986). Pushing the state of the art in CC composites is the piston for internal combustion engines (Miller and Grimes, 1982; Taylor, 1985). The CC piston (Fig. 15.25) would perform the same way as any piston in a reciprocating internal combustion engine while reducing weight and increasing the mechanical and thermal efficiencies of the engine. The CC piston concept features a low piston-to-cylinder wall clearance; this clearance is so low, in fact, that piston rings and skirts are unnecessary. These advantages are made possible by the negligible coefficient of thermal expansion of this kind of CC (0.54 x /OF). (Carbon-carbon composites can have a range of thermal expansion coefficients,
350 Carbon-carbon composites REFERENCES
Fig. 15.25 Carbon+arbon automotive piston.
depending on the processing techniques.) CC material maintains its strength at elevated temperatures allowing the piston to operate at higher temperatures and pressures than those of a comparable metal piston. The high emittance and low thermal conductivity of the CC piston should improve the thermal efficiency of the engine because less heat energy is lost to the piston and cooling system. The elimination of rings reduces friction, thus improving mechanical efficiency. Besides being lighter than conventional pistons, the CC piston can produce cascading effects that could reduce the weight of other reciprocating components such as the crankshaft, connecting rods, flywheels and balances, thus improving specific engine performance (Taylor, 1985). ACKNOWLEDGEMENTS
The author acknowledges Mr. D.M. Curry of NASA Johnson Space Center and H.C. Scott and C.N. Webster of the Vought Corporation for the data, as referenced, on which a portion of the present paper is based. Acknowledgement is also given to Mrs. H.A. Coombs for her valuable contribution in assisting in the formatting of this paper.
Beatty, R.T. and Kipplinger, D.V., 1970, Gas pulse impregnation of graphite with carbon. Nuclear Application and Technology, 8(6):488495. Bokros, J.C., 1969, Deposition, Structure and Properties of Pyrolytic Carbon. Chemistry and Physics of Carbon-A Series of Advances, (ed. Philip L. Walker, Jr.) pp. 1-118. Marcel Dekker, InC. Buch, J.D., 1984, Graphite Crystals - A General Model for Diverse Carbon Forms. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA CP-2357, pp. 119-135. Buckley, J.D., 1967, Statis, Subsonic and Supersonic Oxidation of JT Graphite Composites, NASA TLN D-4231. Cogbum, J.W., Fain, C.C., Edie, D.D. and Leigh, H.D., 1987, Processing C-Shape Pitch-Based Carbon Fibers. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA (2-2482, pp. 185-200. Cook, J.L., F. Lambdin and P.E. Trent, 1970, Discontinuous Carbon/Carbon Composite Fabrication. Carbon Composite Technology - With Special Emphasis on Carbon/Carbon Systems. Proc 10th Ann. Symp. New Mexico Section of ASME and University of New Mexico, pp. 143-171. Curry, D.M., Scott, H.C. and Webster, C.N., 1979, Material Characteristics of Space Shuttle Reinforced Carbon-Carbon. Paper read at the 24th National SAMPE Symposium, 1-9 May, 1979, at San Francisco, CA. Diefendorf, R.J., 1987, Carbon/Graphite Fibers. Engineered Materials Handbook 1: 49-53. Dienes, G.J., 1952, Mechanism for Self-Diffusion in Graphite. Applied Physics 23(11): 1194-1200. Edie, D.D., Fox, N.K., Barnett, B.C. and Fain, C.C., 1986, Melt-Spun Non-Circular Carbon Fibers. Carbon 24(4): 477482. Gray, P.E. and Engle, G.B., 1985, Wettability of Carbon/Carbon Composites and Carbon Fibers by Glass Sealants Used in Oxidation Inhibition. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA 0-2406, pp. 149-162. Johnson, A.C. and Finley, J.W., 1985, Carbodcarbon Composites for Advanced Spacecraft. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA (3-2406, pp. 175-190. Kanter, M.A., 1957, Diffusion of Carbon Atoms in Natural Graphite Crystals. Physics Review 107 (3):655-663.
References 351 Klein, J., 1986, Carbon-Carbon Composites. Advanced Materials and Processes 130 (5):64-68. Kotlensky, W.V. and Pappis, J., 1969, Mechanical Properties of CVD Infiltrated Composites. Proc. 95th Biennial Conf.Carbon Defense Ceramic Information Center, Compilers, pp. 76-80. Lambdin, F. and Cook, J.L., 1971, Fabrication of Carbon-Carbon Composites Electrostatic Fiber Deposition (Flocking). Y-1786 (Contract No. W-7405-eng-26), Y-12 Plant, Union Carbide Corp. Lambdin, F., Cook, J.L. and Marrow, G.B., 1969, Fiber-Reinforced Graphite Composite Fabrication and Evaluation. Doc. Y-1684, "ID4500 (Contract W-7405-eng-26), Nuclear Division, Union Carbide Corp. Maahs, H.G. and Ransone, P.O., 1985, Mechanical Property Evaluation of 2-D Carbon-Carbon Panels Fabricated From a Specialty-Weave Fabric. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA CP-2406, pp. 261-276. Miller, T.J. and Grimes, H.H., 1982, Research on Ultra-High-Temperature Materials-Monolithic Ceramics, Ceramic Matrix Composites and Carbon/Carbon Composites. Advanced Materials Technology, (eds. Charles P. Blankenship and Louis A. Teichman) NASA CP-2251, pp. 275-291. Ohlhorst, C.W. and Ransone, P.O., 1985, Effects of Thermal Cycling on Thermal Expansion and Mechanical Properties of Advanced Carbon-Carbon Composites. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA (3-2406, pp. 289-303. Pierson, H.O., 1968, Development and Properties of Pyrolytic Carbon Felt Composites. Advanced Techniques for Material Investigatioiz and Fabrication 14, National Symposium and Exhibit, Society of Aerospace Material and Process Engineers, Paper II4B-2.
Ransone, P.O. and Maahs, H.G., 1985, Effect of Processing on Microstructure and Mechanical Properties of 3-D Carbon-Carbon. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA CP-2406, pp. 289-303. Ransone, P.O. and Ohlhorst, C.W., 1984, Interlaminar Shear and Out-of-Plane Tensile Properties of Thin 3-D Carbon-Carbon. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA CP-2357, pp. 137-148. Rummler, D.R. and Sawyer, J.W., 1984, Properties and Potential of Advanced Carbon-Carbon for Space Structures. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA (3-2357, pp. 149-170. Sawyer, J.W. and Moses, P.L., 1985, Effect of Holes and Impact Damage on Tensile Strength of TwoDimensional Carbon-Carbon Composites. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA (3-2406, pp. 245-260. Stoller, H.M., Butler, B.L., Theis, J.D. and Lieberman, M.L., 1974, Carbon Fiber Reinforced-Carbon Matrix Composites. Composites: A State of the Art, (eds. J.W. Weeton and E. Scala) Metallurgical Society of the American Institute of Mining, Metallurgical and Petroleum Engineers, Inc., pp. 69-136. Strife, J.R. and Sheehan, J.E., 1988, Ceramic Coatings for Carbon-Carbon Composites. Ceramic Bulletin 67(2):369-374. Taylor, A.H., 1985, Carbon-Carbon Pistons for Internal Combustion Engines. NASA Tech Briefs 9 (4):156-157. Theis, J.D., Jr., Taylor, A.J., Rayner, R.M. and Frye, E.R., 1970, Filament Wound Carbon/Carbon Heatshield SC-11FW-Y12-7, A Process Histoy. SC-DR-70425, Sandia Labs. Webb, R.D., 1985, Oxidation-Resistant Carbon-Carbon Materials. Metal Matrix, Carbon and Ceramic Matrix Composites, (ed. John D. Buckley) NASA CP-2406, pp. 149-162.
HAND LAY-UP AND BAG MOLDING
16
D.R. Sidwell
16.1 INTRODUCTION
Table 16.1 Advantages and disadvantages of hand lay-up structures
This chapter presents practical fabrication and tooling methods that have been successful for Adva n tages high performance applications. Increasing perDesign flexibility formance requirements limit traditional hand Large and complex items can be produced (’wet’) lay-up methods. Today’s prepreg resin Minimum equipment investment is necessary systems allow for longer working life and low Tooling cost is low cure temperatures. Improvements have made The start-up lead time and cost are minimal the current prepreg systems the choice for high Design changes are easily effected Molded-in inserts and structural reinforcements performance composite structures. are possible Composites are a combination of a high Sandwich constructions are possible performance resin matrix and various fiber Proto-typing and pre-production method for combinations, which have brought about a molding processes change in the engineer’s traditional approach Semi-skilled workers are needed, and worker and allow matrix alloying for specific structraining is minimal tural applications. Improved materials, Disadvantages analysis and manufacturing methods have let traditional composite structures reach new The process is labor-intensive limits of achievement. Table 16.1 presents Only one tooled (molded) surface is obtained Quality is related to the skill of the operator some of the advantages and disadvantages of It is a low-volume process hand lay-up of composite structures. Longer cure times required Composite structures have been manufac- The waste factor can be high tured from fiberglass and wood to some of the most expensive 800 GPa pitch fibers for space applications. The engineer’s imagination is the a structural part upon which the external paint only limitation in the fabrication of composite finish is applied. The reverse of this procedure structures. Large, complex co-cured structures occurs with composite materials fabrication, are successfully manufactured with all engi- where the formless materials harden and take neering and manufacturing disciplines the shape of the container into which they are applied. The makeup of the resulting laminate working as one. In the fabrication of wood or metal prod- is an elementary engineering material. The ucts, flat sheet stock is joined together to form physical properties can be changed by varying the resin and fiber ratio, the type and direction of reinforcement and the type of resin matrix. Today bag molded (vacuum and pressure) Handbook of Composites. Edited by S.T. Peters. Published composites provide higher performance and in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Histoy 353 are competitive with alternate types of constructions that result from optimization of process controls, design refinements and improved materials. Composite designs and bag molding processes are interrelated by production goals, the desire to control manufacturing costs and by service or mission requirements. The responsibilities shared by fabricators and designers include: 0
0 0
0
reproducibility and knowledge of processing characteristics of candidate materials; knowledge of requirements for facilities and tooling including provisions for assembly, inspection and quality control; reduction of total part count; avoiding design as replacement for sheet metal; selecting the best material for the application.
16.2 HISTORY
Composite materials first arrived in the New World with the early Spanish explorers who produced adobe bricks for the building trades from a mix of straw and adobe clay. Modern day users developed the chopper gun and fiberglass roving that use low cost inexperienced labor; many fabrications utilize ’wet’ lay-up for making bath tub enclosures and truck hoods. Early use of carbon fiber composites was primarily for secondary structure aerospace applications. In the early 1970s, the first woven graphite fabric was developed for aerospace structural applications. Complex and highly detailed hybrid (unidirectional and woven carbon materials) structures were developed for an intercontinental ballistic missile (ICBM) program. It did not take long to integrate the woven graphite prepregs into mainstream production, reducing labor costs over 70% from that of the previous unidirectional tape structures. During the 1980s, composite hybrids were developed using the best of all material forms available. The hybrid structures utilizing both unidirec-
tional and woven product forms (fiberglass, graphite and aramid) were developed for commercial applications; aerospace structures were the prime users of 100%graphite materials. The fabrication of a 1979 Ford LTD with a finished weight of 1138 kg (2504 lb) proved that graphite could be used as reinforcement for just about any automotive component including bumpers, hoods, frames, doors and drive line. The major drawback to automotive use of these composites was the processing cost. During the 1990s, processing times were reduced to acceptable levels (less than 5 min) for automotive components, using resin transfer and resin injection molding. Improved mechanical properties, lower costs and improved resin systems allowed for greater inroads for fiber-reinforced plastics in the automotive industry. The use of ultra high modulus pitch-based carbon fiber in modern communication satellites has led to refinement in the art of hand lay-up and vacuum bag molding. The primary structural element of some communications satellites is a central cylinder. A typical cylinder is about 180 kg (80 lb) and supporting a 13 000 kg (6000 lb) payload at launch. A typical cylinder consists of 520GPa (75msi) pitch fiber skin laminates and aluminum honeycomb core, the structure is oven co-cured utilizing a nylon vacuum bag. The requirements for ever lighter structures created the need for ultrathin unidirectional prepregs. Solar array panels for a modern communications spacecraft required the use of 0.0004 cm (0.001 in) thick unidirectional 520 GPa (75 msi) graphite prepreg. Solar array panels, made with three-ply skins bonded to an ultralight honeycomb core, are manufactured on a regular production basis, utilizing vacuum bag molding process and oven cure. An example of what can be involved in a high visibility composite materials application was the 1992 America’s Cup Challenge, the first application of graphite composite materials for the America’s Cup racing yachts. The example utilized here is the ’Spirit of
354 Hand lay-up and bag molding Australia’ program entered by the Darling Harbor Yacht Club, Sydney, NSW, Australia. (See section 16.7.2.) Design and analysis were performed by Ian Murray and Associates, yacht designers; the design parameters were established by the America’s Cup Rules Committee. The program objective, determined in large part by available funding, was to manufacture one America’s Cup racing yacht. A program of this magnitude requires a substantial amount of material. The ’Spirit of Australia’ was a relatively small program compared with some of the other syndicates. Twenty-seven America’s Cup yachts were fabricated, representing eight countries, with a total investment of over $600000000. The material combinations utilized in the hull construction of the various entries included Spectra, aramid, graphite and wood. The overall goal was to acheve the specified minimum laminate density and still retain the required load carrying capabilities. Various hybrids of Spectra/graphite, aramid/graphite and graphite/wood were utilized to achieve the design results. The program for ’Spirit of Australia’ required much dedication from supporters and suppliers to make it to the elimination trials with competitors. There were significant other concerns in addition to the normal decision factors in selecting composite materials for a program, such as mechanical properties, tack, resin flow and gel. These other concerns can have a large influence on the program outcome and require planning the materials’ delivery schedule, testing, packaging, export license, custom formalities and shipping method. If the program is within a few hours flight time, the prepreg is usually packed with dry ice and shipped on the next available aircraft. On international programs, the weight of just the dry ice can equal the material cost and delivery takes 3 4 days. Scheduling becomes the critical path in overall planning. In the case of ’Spirit of Australia’, America’s Cup challenge 1993, air shipments were a minimum of three days. The three days were a best effort to get the mater-
ial from the supplier, to the freight forwarder, airline, customs, importer clearing house and then delivered to the manufacturing facility. This effort can go astray; the shipment can be off-loaded at an intermediate stop or offloaded at the incorrect destination. Sea shipping requires 30 days to get material from the manufacturer to a sea forwarder, providing the container is off-loaded as scheduled. All of these considerations combined with availability of funds and tight fabrication schedule play a primary role in design and fabrication decisions. 16.3 MATERIALS
Fiber reinforced composite materials are now being used to satisfy the high strength and/or stiffness-to-weight requirements of aerospace structures. Fibers of extremely high strength are incorporated into a relatively weak matrix that is used to bind the fibers; the combination results in a light, extremely strong structural material. Experience and theoretical studies have shown that fibers will always be stronger than bulk materials. Some of the properties that can be improved by combining the constituents into a composite are strength, stiffness, wear resistance, corrosion resistance, fatigue life, temperature resistance and electrical and thermal conductivity. The relationship between resin viscosity and the cure cycle can be used to obtain maximum performance in a composite structure. Figure 16.1 shows the relationship between resin viscosity and a stepped autoclave cure cycle. The curves shown are for a typical 121°C (250°F) curing epoxy resin system. The viscosity will vary with a change in heat-up rate and temperature. This information relating the effect of rate and temperature is available from the material suppliers. Using this information as a starting point, a cure cycle can be defined for a specific composite structure. The cure cycle production capability becomes increasingly important as the complexity of the composite structure increases. To help in planning a
Materials 355 Apply 689 KPa (loo PSI) when Part Reaches 78OC +3"C, Hold 15 Heat at 2% per Minute Minutes
Heat at 2°C Per Minute
Curecycle
I
Hold 120 Minutes at 120°C
II
1
I
Down to 75 "C II
I
125
Resin Flow-
c .1
0
201
101
401
301
60I
501
200
Time (Minutes)
Resin Flow
1
I
I
I
Initial Resin Flow
I I
I
Removal of Entrapped Air and Resin Bleed
Maximum Resin FIOW
Resin Gel
I I
b
Resin Crosslinking and Cool Down
Fig. 16.1 Typical autoclave stepped cure cycle.
composite structure, the material supplier can spray-up. In contact molding, after a gel coatprovide information on specific applications, ing is applied to the mold (if required), processing parameters, material properties, fiberglass chopped mat, cloth or woven roving test data and fabrication methodology for the is placed into the mold, saturated with resin material. Because most materials have compet- and brushed or rolled to compact the material itive products available, one must consider not and remove entrapped air. This process can only the price and delivery available but also also utilize 'wet' prepregs. The wet composite local technical support in selecting a supplier. is then vacuum bagged to compact the materThere are two basic types of composite ial and remove entrapped air. Spray-up matrix fiber combinations: chopped-fiber reinforced composites consist of relatively reinforced composites and continuous-fiber short fibers mixed with a matrix or binder reinforced composites. The selection of the material resulting in a composite with generproper matrix material for use in a particular ally lower physical properties. Continuous fibers in a matrix material result structure is influenced by many factors including weight limitations, environmental in a composite that exhibits higher strength and stiffness than the chopped-fiber counterparts. resistance and types of loading. 'Wet' lay-up procedures are divided into Fabrication of parts that use continuous-fiber two main types, contact or open mold and reinforced composites can be achieved by
356 Hand lay-up and bag molding several techniques such as hand or machine 1. Thermoplastics. Thermoplastics may be lay-up, filament winding and pulhusion. shaped, remelted and reshaped, much like candle wax. Examples of thermoplastics are Hand lay-up composites are a special case nylon, polysulfone, polyethylene, polyvinyl of continuous fiber reinforced composites. Layers of unidirectional or woven bidirecchloride and polyarylsulfone. tional composites are combined to result in a 2. Thermosetting. Thermosets, once shaped and material exhibiting desirable properties in one set, cannot be remelted and reset (like a or more directions. Each layer is oriented to cooked egg it cannot be uncooked). When achieve the maximum utilization of its properheated, a thermosetting plastic becomes ties. Layers of different materials (different increasingly infusible. It undergoes an irrefibers in different directions) can be combined versible chemical change called polymerization (a process in whch simple to further enhance the overall performance of molecules combine to form more complex the laminated composite material. Some of the initial considerations that must be considered ones) during curing of the material. before undertaking the manufacture of a comExamples of thermosets are polyesters, posite structure are: epoxies, phenolics, cyanates and polyimides. 0 performance goal to be achieved; 0 can the structure be designed to reduce part In advanced composite materials manufaccount? ture, there are controls to provide a workable 0 do not duplicate a sheet metal design; resin system. The main control points include 0 manufacturability - in-house or outside volatile content to minimize voids and delamsupplier; inations, tack that controls the handling of the 0 overall cost compared with performance; prepregs, flow to allow removal of excess 0 use only the amount of expensive materials resin and entrapped air and gel that deterneeded. mines the point at which the resin starts cross linking. Many structures can be designed and manufactured with fiberglass laminates using small amounts of graphite, aramid or Spectra, Volatile content controls hybridized to obtain the additional stiffness or reduced density. The high performance of Quality control test results do not distinguish graphite (both PAN and Pitch) is attractive inert from reactive volatiles but are used to for fishing rods, tennis rackets, golf clubs and establish volatile content values. They provide racing applications. Special incentives for qualitative evaluations of prepreg advanceperformance can make cost a secondary ment, degree of volatilization of solvents and amount of degradation due to aging. Volatile objective. contents measured prior to lay-up are compared to values when the prepreg was first 16.3.1 MATRIX MATERIALS received and may provide indications of Plastics are used as the matrix material in excessive aging. most fiber reinforced composites. These materials are generally of high molecular Tack controls weight and intricate molecular complexity. Their form and properties can be transformed Tack is the adhesion characteristic that is conby the techniques of applying heat and/or trolled to facilitate lay-up operations. It is pressure. In general, plastics can be divided affected by resin and inert volatile content, into two categories: prepreg advancement and humidity and
Materials 357 temperature of the lay-up room. Sometimes, tack is increased by increasing resin and volatile content, less prepreg advancement, or a slight increase in lay-up room temperature. At other times, tack can be modified by the prepreg manufacturer by changes in resin formulations or new additives blended into the resin. Alterations to the formulations should not be accepted without prior requalifications by the user. Often, properties are severely affected for a cured laminate; therefore, performance and durability must be verified. Prepregs with excessivelyheavy tack generally cannot be handled without grossly disrupting resin distribution and fiber orientation or causing a roping (fiber bundling) of the composite laminate. Excessive tack makes it difficult to achieve reproducibility in composite structures since undetermined amounts of resin are always removed when the release film or backing is separated from the prepreg. In general, all the disadvantages of wet lay-up systems are inherent in the excessively tacky prepregs. Prepregs with no tack are either excessively advanced or have exceeded their normal storage life. Such materials cannot attain adequate cured properties and should be discarded. Exceptions are silicones and some polyimides that can only be prepared with no tack. Layups with these materials are limited to those occasions where lower mechanical properties can be tolerated in exchange for improved heat resistance or electrical properties. Thermoplastic prepregs typically do not exhibit much tack. Most prepregs are staged to attain more manageable tack. The tack qualities should be adequate for adhering the prepreg to prepared molding surfaces or preceding plies for a layup and still light enough to release the backing film without loss of resin. Provided that cured laminates will not be adversely affected, tack qualities can be specified to require the prepreg to remain adhered to the backing until a predetermined force is applied to peel it off. Tack requirements can be modified to suit
individual plant fabrication conditions. Air conditioning minimizes local temperature and humidity sensitivities; reducing the temperature makes heavy tack manageable. Judicious use of hot air guns can improve dry tack. Flow controls Flow measurements indicate resin capabilities to fuse successive plies in a laminate and bleed out the void producing gas reaction by-products. Flow is also an indicator of prepreg age or advancement. It is often desirable to optimize resin content and tack to attain adequate flow but, in some cases, flow is controlled by thickening additives in a resin. Gel times Gel time is an indicator of the degree of cure of resin systems. The amount of staging or advancement limits gel time for prepregs. Most prepregs are formulated to attain a useful life (out time) of 8 days or more at standard atmospheric conditions. Out time can be prolonged by storage at -18°C (O'F), but each time the prepreg is brought to equilibrium at room temperature, the useful life is shortened. Gel time measurements can be used as quality control verifications. Criteria based on those results determine whether to initiate more costly property testing or to dispose of an overage prepreg. Storage Most advanced composite hand lay-up materials that are utilized today are time/temperature critical. Precautions must be taken in handling, storage and processing. During the initial processing, where resin is impregnated onto the reinforcement fiber, the degrading of the materials starts the moment the resin is applied to the reinforcement. However, cases exist where material has been stored at -18°C (0°F) for as long as 5 years and was capable of molding a commercially
358 Hand lay-up and bag molding acceptable laminate. A new generation resin fiberglass fiber system, the cyanates, tends to be more stable 0 high impact strengths; than epoxies and outlife has been increased 0 good overall performance; from 10-14 days to 3 0 4 5 days. Storage at 0 low costs; -18°C (0°F) has also increased from 6 months 0 available in many forms; 0 higher density. to 1year. In planning a project, one should consider Aramid fiber the amount of material needed within a given 0 high impact strengths; period of out time to complete the lay-up and 0 high specific modulus; cure of a composite component. If one can fab0 poor compression strengths; ricate enough composite parts in a 10 day 0 high costs; period to utilize a 22 kg (50 lb) roll of mater0 difficult to handle. ial, then that is the size that one should procure. If not, reduce the size of the roll Spectra fiber accordingly. By doing this, the material roll 0 very low density; has to be removed from the freezer only once 0 difficult to handle; and the recorded out time can be accurately 0 bond interface poor; maintained. 0 very high impact resistance; 0 high costs. 16.3.2 REINFORCEMENTS
In continuously reinforced composites, the fibers carry nearly all of the load applied to the system, with the matrix transferring the load into the fibers. In chopped fiber reinforced plastics, the load is distributed between the fibers and matrix, depending upon the nature of the constituents,percentages of each and the orientations of the fibers. Selection of the proper fiber for a specific application depends upon the type and magnitude of loading, operating temperature and cost limitations. The following are attributes of several applicable fibers:
Carbon fiber reinforced composites Carbon/graphite fiber is an attractive material because it has the best balance of material characteristics to compete with metals. Carbon fiber is the end product of a series of manufacturing processes that begin with the basic PAIV precursor that is carbonized and graphitized. It is then collimated, woven or, chopped to become tape, cloth, or molding compound when prepreged with an appropriate matrix material. Carbon composites offer many advantages over conventional materials. Major advantages are: 0
PAN-based carbon fiber 0 high strengths; 0 improved strain to failure; 0 workable; 0 available in many forms. Pitch-based carbon fiber 0 high to ultra high modulus; 0 low strain to failure; 0 high cost; 0 can be difficult to handle.
0
0 0 0 0 0
high specific strength and stiffness; can be woven into cloth; dimensional stability at temperature; tight product control; ease of manufacturing; high fatigue resistance; extremely complex structures can be fabricated.
While offering significant advantages, carbon fiber does present the following problems relative to conventional materials that must be
Materials 359 considered when designing for specific applications: low strain to failure compared to most metals; relatively low impact resistance; potential for galvanic corrosion; potential matrix bonding problems; low interlaminar shear; high cost; chemical resistance.
Non-woven and collimated tape During lay-up with unidirectional tape, there are several techniques that will help achieve a cost effective structure. As soon as unidirectional tape has been selected, start preliminary planning as to ply orientations and develop basic patterns for the initial fabrication. The results of this preplanning will be that wastage will be minimized and better quality can be built into the finished part, Using unidirectional tape on contoured surfaces presents a number of lay-up problems. As the tape is placed on the mold surface, the orientation can change dramatically during its application. In order to minimize angle changes, narrow pieces of unidirectional tape can be used. Scrap material can be used for local reinforcements and to fill gaps between pieces of unidirectional tape. Gay filling pieces up to 3.2mm (1/8 in) wide are generally acceptable,
Woven fabric Joints with woven materials can be more of a challenge than those with unidirectional tape. On a highly stressed composite structure, the selvage edge must be removed to minimize the discontinuities in the lay-up joints. Joints in succeeding plies must not be adjacent but staggered by a minimum of 51 mm ( 2 in). Hybrid materials
Many combinations of composite materials have been developed over the past 15 years.
Some are better than others; some have application only to specific needs of a particular structure. Some of these are presented to show the basic selection process that is required in the initial design phase and the manufacturability of each product form into a successful composite structure. All fibers can be obtained in many different forms: unidirectional tapes, Unidirectional tows or yarns, woven cloth, chopped and in hybrid Combinations. The combinations are unlimited in hybridizing fibers together. The following list shows some of the hybrid relationships and Table 16.2 shows some of the properties of composite material forms.
Carbon/fiberglass high impact strengths; 0 high modulus; 0 high performance; 0 cost effective. 0
Spectralcarbon improved bond interface; 0 little improvement in overall properties; 0 low density. 0
Aramid,@ergIms 0 reduced overall costs; 0 limited applications. Aramid/graphite 0 limited applications; 0 high costs; 0 highly resistant to cracking; 0 better impact resistance. 16.3.3 EXPENDABLE MATERIALS
Supporting materials can contribute to a program’s success as much as the composite structures. If there are problems with the vacuum bag, release film, or sealant tape, the final product quality can be compromised although the overall unit cost of most of these materials is low. Specifications to control support materials have been slow to develop with the industry. There have been problems with expendables such as release coating of porous
360 Hand lay-up and bag molding Table 16.2 Limits of composite material forms
Material form
Fiber modulus GPa (mi)
Tensiltf properfies MPa (ksi)
Compressionc properties MPa (ksi)
Tensile modulus GPa (mi)
Fabric
Graphite 8HS Graphite 5HS” Graphite 5HSb E-glass 8HS Kevlar‘ 49 8HS
228 (33) 228 (33) 537 (75) 72.4 (10.5) 130.9 (19.0)
586-724 (85-105) 758-896 (110-1 30) 927-1114 (132-158) 462483 (67-70) 510-579 (74-84)
522489 (80-1 00) 758-827 (110-120) 240 (34) 469-510 (72-74) 138-1 79 (20-26)
62-83 (9-12) 76 (11) 188 (26.7) 22.0-26.2 (3.2-3.8) 29.6-33.8 (4.34.9)
228 (33) 228 (33) 345 (50) 520 (75) 720 (105) 780 (114) 72.4 (10.5) 86.8 (12.6) 130.9 (19)
1586-1 792 (230-260) 1930-2137 (280-310) 1069-11 72 (155-1 700) 1760 (256) 1760 (256) 1760 (256) 1034-1171 (150-170) 1241-1379 (180-200) 882 (128)
1517-1 724 (220-250) 1379-1 723 (200-250) 896-1034 (130-150) 420 (61) 330 (48) 330 (48) 483-620 (70-90) 552-690 (80-100)
124-138 (18-20) 138-152 (18-20) 200-228 (29-33) 300 (44) 420 (61) 450 (65) 41-55 (6-8) 48-62 (7-9) 64.9 (9.4)
Unidirectional tape
Graphite Graphite” Graphite Graphiteb Graphite Graphite
E-glass S-2 glass
Kevlar‘ 49
a
-
High strength >3447 MPa (>500ksi) PAN based graphite fiber High strength >3630 GPa (>SO0ksi) pitch based graphite fiber Composite laminate properties
Teflon fabric where there was insufficient coating material to provide an effective release, resulting in a bond between the Teflon fabric and the composite structure and when using a perforated release film, the perforations were torn rather than pin pricked, allowing excessive resin bleed. Specifications would help limit the inconsistency of some of these products. If one
encounters a problem utilizing support materials, contact the supplier who can provide technical assistance on its products. During the lay-up of composite structures, care must be taken to insure that all areas are covered with a release or separator film. The bleeder and/or breather will bond nicely to composite laminates if there is no separator.
Tooling 361 16.4 TOOLING
for steel, include superior durability, ample tolerance for elevated temperature service and Tooling includes materials, equipment, or good thermal conductivity. forms onto which (or into which) the product Ceramics have favorable characteristics for is made, assembled, or cast. Tooling issues are molds. They have the lowest coefficient of a result of a number of interacting requirethermal expansion and their thermal conducments that are considered when selecting the tivity approaches that of some hardened tool most cost effective tooling. These requiresteels. However, ceramics are brittle at ambiments are as follows: ent temperatures and they must be protected ability to achieve a uniform heat up rate, from processing and handling hazards. One taking into account the mass of the tool; way to get protection is to enclose ceramic allowing sufficient movement of the lami- inserts in a steel case. nate while achieving pressure and Dual steel molds are candidate tools for compaction in all regions; reproducing high quality composites. accommodating resin flow; However, these tools are costly and producfacilitating or allowing for removal of cured tion quantities are often not sufficient to part; amortize tooling costs at competitive prices for realistic tolerance for tool and laminate; the production items so less costly alternatives finishing requirements - coarse finish can are desirable. create lock-on problems; Aluminum molds are less costly. Although adequate area for applying sealant tape for thermal conduction is better for aluminum vacuum bagging and test coupons as than for steel, the tools are less durable and the required; thermal expansion is large. Shallow or flat provisions for vacuum fittings and han- mold plates are usually limited to cures below dling of tool. 177°C (350°F). Other metal tools include sprayed or electroformed molds reinforced Tooling is less expensive for vacuum bag and with cast backings. Alternate types of tooling autoclave/oven molding than for matched die can include composite molds usually based on molding methods. Molds and molding plates high temperature resistant cast or laminated are required to withstand curing conditions epoxy resins. without distorting or degrading and to tolerMaster forms for laying-up composite tools ate handling during the fabrication processes. can also be fabricated using any of the materiThey are not necessarily resistant to unbalals described; a mock-up model of the item anced pressures. The higher costs of composite may be used, or plaster masters can be pretooling can be amortized by taking advantage pared from models. The quality of the plaster of the improved capabilities to mold complex masters depends on the strain compatibility constructions. Composites that may ordinarily between the plaster and its reinforcements and require secondary bonding are often more ecoon the condition of the hardened surface. nomically co-cured. Composite tools can be laid-up using fiber orientations that most closely match the expansion of the items to be produced. 16.4.1 MOLD PREPARATIONS Fiberglass and graphite fibers are the principal Coefficients of thermal expansion for conven- reinforcements. Woven fiber reinforcements tional tooling materials and composites are are the most economical to use. Mold maintegiven in Chapter 25. For metals, the coefficient nance is best relegated to specialized for steel compares most closely with the coeffi- personnel while preparations for the bag cients for the composites. Other characteristics molding processes are assigned to production
362 Hand lay-up and bag riiolding personnel. A successful practice is to provide production personnel with soft tools and solvents that do not degrade the molding surfaces. If the production tools and solvents are inadequate for removing debris and cleaning, the molds are taken out of service for maintenance, repair, or replacement. After they are returned to service, they are solventwiped clean and mold release agents are applied. 16.4.2 RELEASE AGENTS
Release agents for bag molding composites include carnauba paste wax, aerosol dispensed compositions that contain carnauba, fluorocarbon resins, or silicone resins, plastic films and metal foils. On most occasions, the wax or resin mold releases do not contaminate the composite surfaces excessively nor prevent subsequent secondary bonding or coating operations. Prior to bonding, the composite surfaces are cleaned with solvent and lightly sanded to remove resin gloss. The user should be aware of national or regional limitations on solvent usage due to toxicity or ODS (ozone depleting substances) concerns. Peel plies can be used to protect clean surfaces for primary adhesive bonds. Plastic films, metal foils and sprayed metal coatings also serve as release agents when they are integrally laminated to the co-cured composites. Both the polished wax surfaces and the sprayed wax coatings are excellent mold releases for composites cured below 121°C (250°F). However, the wax degrades and discolors the composites at higher molding temperatures. Commercial fluorocarbon mold releases are used for higher cure temperatures. FEP (fluoroethylene propylene) mold releases form a continuous film on mold surfaces. Although the condition of the release film is easy to maintain below 177°C (350"F), the coating degrades at higher temperatures above this. Fluorine which is noxious, corrosive and highly toxic can be released from the polymer above 177°C (350°F).
PTFE (polytetrafluoroethylene) is stable and is often contained in mold releases for service in excess of 260°C (500°F).The mold releases contain suspensions of micropulverized PTFE in a volatile dispersant. Depositions on mold surfaces do not form continuous films, but the PTFE particles provide excellent dry lubrication for the release of the cured composite. Furthermore, the residual particles on the cured composite surfaces are easily removed with a solvent-wipe. Since a variety of commercial mold releases that contain fluorocarbon (or equivalent)are on the market, it is essential that manufacturers' recommendations on uses and limitations be scrupulously followed. Silicone oils and greases are to be avoided since they are the most persistent contaminants of molded composite surfaces. They release secondary bonds and coatings from composite surfaces as effectively as they release the cured composites from the molds. Silicone oils and greases migrate and defeat most removal attempts. They contaminate the solvent wetted cloths and sand papers so that instead of removing the silicones, they spread them. Contaminated surfaces may be salvaged for painting by sandblasting with virgin grit. Table 16.3 summarizes the precautions necessary for successful hand lay-up and vacuum bagging operations 16.4.3 MOLD DESIGN
hlold design is a function of cost and projected life and/or use. A production mold should be made carefully from the best materials. Such a mold will be designed by an experienced designer who will incorporate the necessary thicknesses, materials, structural reinforcement and hardware required for the intended use. 16.4.4 PATTERNS (PLUGS)
A pattern (plug) is a temporary form in the exact shape, contour and finish of that to be molded. (If the outer shape of the items desired the inside contour is used.) Patterns
Tooling 363 Table 16.3 Hand lay-up and vacuum bagging precautions Mold release application Selection of correct mold release ‘Seasoning’ new tool to insure coverage a minimum of three coast Compatibility of resin system and mold release Repeated applications can cause excessive buildup Sealant tape Ease of use, release from backing Double or single strip application Removal after cure Cheapest when suitable Flash breaker tape Check compatibility Useful for holding prepreg during lay-up and heat debulks Separator film Compatability and will release as intended Quality control on perforations Drapability for complex shapes Elongation (%) (high elongation) Bleeder Not all bleeders the same, select for application Ensure no compaction during cure Potential seal off during cure Excessive bleed can saturate bleeder
are made of many materials: wood, plaster, plaster/metal and other combinations. Almost any material can be considered as a pattern material if it holds its shape. It is assumed that when only one large composite structure is required, such as a 23 m (75ft) America’s cup yacht hull, the cost of making a pattern and a mold in order to make the hull may not be justified. However, the construction of a pattern that becomes a male plug can be cost effective for high performance composite structures. To avoid excessive cost with this tooling approach, one must remember that it is the total cost of the end product not just the cost of the pattern (plug) that must be considered. A limited use plug for a large marine hull 14-23 m (45-75 ft) in length would be made as follows: a simple wood frame to
Breather Able to malntain vacuum path Care not to puncture nylon film Bagging film Higher the percent elongation the more forgiving Pliability Defect free Select for temperature performance Thermocouples Through bag/sealant tape, potential vacuum leaks Placed on outside of bag - reliability Vacuum fittings Caution not to allow resin to fill 0 Source for potential leaks Removal from bagging film before disposal 0 Integral with tooling Vacuum lines Ensure fittings do not leak Hose has not been crushed 0 Not pinched off Not filled with resin
the inside contour of the hull is constructed from stanchions and stringers and covered with strips of wood, with a laminate of 2 45”, 90” layers. The tooling plug is finished and covered with mold release. 16.4.5 INTEGRATING INSPECTION AND MACHINING
A machine tool is for machining. If it isn’t making chips, it’s wasting time, so keep non-cutting time to an absolute minimum. That’s the standard philosophy most shops try to live by. In fact, many shops are investing in pallet shuttles, quick-change fixtures, tooling systems, rapid transverse fixtures, programming and scheduling systems to keep spindles turning and cutting tools w o r h g at optimum capacity.
364 Hand lay-up and bag molding
Inspecting the workpieces right on the fiveaxis mills, has the equipment functioning as both machine tool and coordinate measuring machine, so that the production of parts and producing inspection data become equally important. Such a radically different plan means that design, numerical control (NC) programming, machining and inspection cannot be separate functions.Just as each machine tool would have to serve more than one role, one computerized database would have to share the same information with designers, programmers, operators and inspectors. This combination allows a shop to machine, inspect and analyze any surface without removing the work piece from the five-axis machine. This system helps produce higher quality tooling with significant gains in productivity. A large machine bed will accommodate unusually long workpieces and also leave room for smaller workpieces to be clamped on one end while another workpiece is being machined at the other end (see Fig. 16.2).
parts. If the workpiece will not fit into the hard gauge, it has out-of-tolerance features and will not fit mating parts. The workpiece is rejected. A soft gauge can be used to make similar either-it-fits-or-it-doesn’t comparisons. Instead of placing two physical objects together, two CAD models are laid one over the other on the graphics screen. The software version of the checking fixture is the soft gauge. The software version of the workpiece to be inspected is a geometric model constructed from inspection data. Out-of-tolerance conditions will be just as conspicuous in this comparison, but analysis is far more complete and much faster. Moreover, a soft gauge is created directly from the original design data. Because it is created on a computer screen instead of in a tool room, a soft gauge can be constructed quickly and modified easily. It spares the high cost of building and validating a hard gauge.
Closed-loop machining
Closed-loop machining begins with electronic data representing part geometry from the cusSoft gauge tomer. This data describes the outer surface of A soft gauge can be compared and contrasted the customer’s end product. The CAD system with a ’hard’ gauge such as a conventional then creates a 3D model of its surface. Once checking fixture used for inspection and qual- this surface has been established, all manufacity control. If a workpiece drops into the hard turing operations will be derived from and gauge, it is acceptable and will fit with mating related to it. NC tool paths will be generated
Fig. 16.2 Closed-loop machining - mounting various tools. (Courtesy of Coast Composites, Inc.)
Fig. 16.3 Closed-loop machining - touch probe inspection tool. (Courtesy of Coast Composites, Inc.)
Tooling 365 from it. Using dynamic display of the tool path, programmers can visually verify the NC program, check clearances and make sure gouges are avoided. The inspection path will be generated from the same surface geometry. By referencing the soft gauge, the inspection path will be sure to include checks of all critical features. The path of the probe can be visually verified in the same way as the NC program. After executing the NC program, the workpiece can be inspected immediately using the touch probe in the spindle (see Fig. 16.3). This inspection can be considered in-process, because the workpiece is still fixed on the machine tool and can be remachined without being moved or refixtured. This approach is called closed-loop machining. Results of this inspection routine are automatically used to create a 3D model of the features checked. By comparing this model to the soft gauge, any out-of-tolerance conditions can be identified. It will also show where additional machining passes will be required. Final inspection can performed on the machine tool. These results are compared to the soft gauge again to verify that the contoured surface of the graphite tooling will produce the intended part. Using this machining approach for inspection on the machine tool reduces inspection time by 80%. The biggest savings come from eliminating workpiece moves and additional setups and from streamlined programming of the inspection routines. By integrating inspection and machining, overall manufacturing cycle time can be reduced by 30%.
the composite materials. For complex shapes with integral stiffeners, each block of silicone rubber is wrapped on all but one side, the side in which the tooling rubber is removed (see Figs 16.4(c) and (d)). In tooling a thermal expansion molding, it is best to avoid using rubber on both sides of a laminate as illustrated in Fig. 16.4(e)unless straight edges are not critical. Vacuum bag assist (see Fig. 16.4(f)) provides an alternate method. The linear thermal coefficient of most silicone rubbers that have been measured fall into the range of 1-2.1 x This range is consistent over a 23-246°C (75480°F) temperature range. The rubbers are said to have a linear expansion of approximately 17 times that of carbon steel which is why they are used to mold composites by thermal expansion molding techniques. Precautions in mixing some silicone rubber compounds are required if full potential is to be achieved. During prolonged storage, the catalyst tends to separate and settle to the bottom of the container. Mixing the catalyst prior to adding it to the base rubber will allow correct mixtures and long tool life.
(a) Compression molding
16.4.6 THERMAL EXPANSION MOLDING
Thermal expansion molding techniques are utilized for special applications of small complex composite structures and composite tubing with critical outside surfaces. Figures 16.4(a)and (b) illustrate the methods allowing the expansion of the silicone rubber to provide the required pressure for the compacting of
(b) Oven cure critical outer surfaces
Fig. 16.4 Thermal expansion molding. (Continued on next page.)
366
Hand lay-up and bag molding
(c) Enclosed molding
Incorrect mixtures will be light or dark in color and materials such as Silastic J@ will start to crumble within a few thermal cycles. The use of thermal expansion rubber can be hazardous if not planned well. The tooling rubber can exert up to 6.9 MPa (1OOOpsi) during its first few thermal cycles. The tooling rubber requires the minimum of five full cure heat cycles free standing after the initial cure in order to stabilize the expansion characteristics. The tooling is capable of producing 50-75 composite parts before having to be replaced. One other problem with thermal expansion rubber is with its removal from the composite structure. Sharp pointed objects will have a very lasting effect on tool life; once the rubber is damaged, it will continue to tear, needing replacement much sooner than usually required. Silicone rubber is very slow to cool down and extra time must be allowed because the rubber is impossible to remove from the composite part until it has shrunk back to its original size. Putting the thermal expansion rubber tooling into a freezer can accelerate the manufacturing cycle. However, since some of the tools are heavy, due to steel outer encasements, a 12-15 h cooling down period should be planned into the manufacturing cycle. Experimentation is suggested with this molding method since extreme pressures can be generated and undesirable results may occur if the molding method is picked arbitrarily.
3
~~~~
(d)Negative draft molding
r - - - - 1
I
(e) Oven/press cure
1
16.5 BAG MOLDING PROCESS
(f) Vacuum bag assist
Silicone rubber Steel molds
IICompositelaminate Fig. 16.4 Continued.
Molding methods include vacuum bag, pressure bag, oven and autoclave molding. Bags, the thin and flexible membranes or silicone rubber shapes, separate the laid-up constructions from atmospheric pressure during composite cures. The bagged lay-ups in autoclaves are usually vented to pressures lower than those applied to the bag. Consolidations and densifications of the lay-ups are achieved by the resulting pressure differentials across the bag contents. Consolidations are achieved
Bag molding process 367 when the separate plies of prepreg in the layups and other adherents are bonded together. Densifications result in reduction of voids and removal of excess resin. Other results desired of bag molding methods during cure include prevention of blistering in the composites, increased controls of pressure and heat application and control of the fiber/resin ratio. Consolidations and densifications of vacuum bag moldings can be achieved by atmospheric pressure alone as the bagged layups are evacuated throughout the cure cycles. The pressure-bagged and autoclaved-cured composites are pressurized by hot gases. Vents to the atmosphere or vacuum provide for the escape of the volatilized reaction by-products and the entrapped air from the curing composites. If the pressures within the bag are not reduced from those applied to the bag, the membrane remains inert and there is no compaction. Of the three methods, vacuum bag molding is least limited as to the size of constructions that can be processed. On a few occasions, 'wet' lay-up vacuum bag molded composites are room-temperature cured. Most are thermally cured to produce improved properties. Thermal cures are best attained in air circulating ovens/autoclaves, but can also be achieved in infrared heated and passive type convection ovens. Pressure bag molding methods are efficient for producing both deeply contoured structures and shallow composites. Sonar domes, radomes and antenna housings are examples of deeply contoured composites.Architectural panels, door panels and aircraft fairings are examples of shallow composites. Heavy molds are built to reproduce deeply contoured structures. Each specialized mold is constructed to withstand the elevated temperatures and increased pressures required for the cures. Shallow items may often be bag molded in modified compression presses. The lower press platens contain vents and vacuum lines. The upper press platens are made hollow to enclose the mold plates together with the laid
up assemblies. When the presses are closed, the sealed chambers are pressurized and heated to attain molding conditions similar to those of an autoclave. Unlike the specialized pressure bag molds, the modified presses are used to cure many different composite constructions. Autoclave and pressure bag molding conditions to 177°C (350°F) and 1379 kPa (200 psi) are routinely attained. Newer, customized autoclaves attain cure conditions that exceed 260°C (500°F) and 3447 kPa (500 psi). Fire hazards are greatly increased at elevated temperatures and pressures. Pressure vessel fires are minimized by uses of fire retardant processing materials and inert pressurizing gas. Fire prevention measures include uses of silicone rubber, nylon or Tedlar bags. Before cure cycles are initiated, the pressure vessels are purged of all enclosed air. After the thermal cure is completed, the pressure vessels and their contents are cooled to 68°C (150'F) before the pressure is relieved and the autoclave is opened. 16.5.1 EXPENDABLE VACUUM BAGGING TECHNIQUES
Bleed-out systems are devised to maintain reduced pressures within the bags' contents. The bagged lay-up includes the bleed out system designed for the composite part. Bagged lay-ups can be bled in two ways: vertically or edge bled. The classical differences between the two can be seen by comparing Figs 16.5 and 16.6. Many of today's resin systems are mostly 'net resin' and do not require any resin bleed during cure. This allows for better control of the resin content of the composite structures. If a resin bleed sequence is preferred, the following sequence can be used as a guide. 0 0
The surface of the mold is prepared with the release agent. The composite plies are applied and rubbed out to remove entrapped air.
368 Hand lay-up and bag molding
I ATMOSPHERIC OR VACUUM VENT
Fig. 16.5 Vacuum bag edge bleeder - schematic. CAUL PLATF
RAG
NOTE HEAVY PROTECTION AGAINST BAG PERFORATIONS DUE TO INCREASED AUTOCLAVE PRESSURES
Fig. 16.6 Vacuum bag vertical bleeder - schematic.
Bag molding process 369 A perforated release film is applied over the composite laminate and extended approximately 3.2 mm (1.25 in) beyond all edges. A predetermined number of bleeder plies are applied over the release film and extended to the perimeter of the lay-up. A perforated release film is applied over the bleeders and extended 3.2 cm (1.25 in) from edge. One or two layers of a non-woven breather is placed over the lay-up and extended over the release film. Sealant tape is applied around the perimeter of the bleeder. The vacuum bag is positioned and sealed. The contents are evacuated and smoothed and the bag is checked and sealed against leaks. The bagged lay-up is ready to be cured.
In any bagging sequence, the types of release film, bleeder, breather and bagging materials used vary from company to company and from supplier to supplier. Each supplier has typical data sheets on expendable materials to acheve the most efficient use of the materials.
f
In a typical vacuum bag lay-up, there are several methods available; some use double sealant tape side/side, some single and some one on top of the other. The best system is the one that works. The side/side method is used to provide some insurance during cure that the bag will not shrink, pulling an edge off, causing loss of vacuum. The over/under method is used to provide ease of placement of ears to allow some movement of the vacuum bag. No matter which method is chosen, it is important to remember that vacuum bags tend to pull down more than expected and can puncture, if bagging is over a sharp object. During the application of a vacuum bag, 'ears' are required to facilitate the uniform application of vacuum to the composite laminate. Vacuum bag bridging is one of the leading causes of resin rich and excessive voids in corners of composite laminates. Figure 16.7 illustrates this common problem. One method of eliminating bridging of the vacuum bag is presented in Fig. 16.8by means of 'ears' in the bag. Another method to help reduce resin rich
SEALANT TAPE
VACUUM BAG
LAMINATE BEING FORMED
Fig. 16.7 Vacuum bag bridging (Morena, J., Advanced Composite Moldmaking; New York, Van Nostrand
Reinhold, 1988).
370 Hand lay-up and bag molding
fSEAMNT TAPE
VACUUM BAG
4- INCH HIGH PLEAT OR FOLD
LAMINATE BEING FORMED
Fig. 16.8 Elimination of vacuum bag bridging (Morena,J., Advanced Composite Moldmaking; New York, Van Nostrand Reinhold, 1988).
Fig. 16.9 Large vacuum bagged structures. (Courtesy of Richmond Aircraft Products.)
Bag molding process 371 and excessive voids in corners is the place- 16.5.2 REUSABLE VACUUM BAGGING ment of an intensifier over the area, usually TECHNIQUES placed between the separator film and There are material and recurring labor cost breather. The intensifier can be molded rubber disadvantages to the use of expendable vacin the radius desired or some sealant tape to uum bags of plastic films for fabricating fill the corner. 'Ears' may be required in sev- production composites. Expendable bags, laid eral sections of a complex part. Experience will up of plastic films and associated sealants, also determine the height of the ear for a specific incur recurring costs. Expendable bags can be application; 10 cm (4 in) is a good starting laid up only once because of degradation durpoint. Some will be smaller and some will be ing handling and the thermal cures. larger depending upon the complexity of the Use of silicone rubber reusable bags can component being vacuum bagged. reduce fabrication costs and defective comThere is essentially no limitation on the size posite parts because of resulting work of thermoset composite structures. The use of simplification and more positive control of the the thermoset vacuum-bagged composites bag molding cure conditions. Figure 16.11 (autoclave, oven or integrally heated) will con- illustrates an example of a component being tinue to provide excellent composite manufactured utilizing a reusable silicone structures for many years to come. Figures 16.9 and 16.10 illustrate some more complex uses for expendable vacuum bags.
b
I
I
Fig. 16.10 Complex vacuum bagging. (Courtesy of Richmond Aircraft Products.)
Fig. 16.11 Disposable vacuum bagging (top); reusable vacuum bagging (bottom). (Courtesy of The Darner Corporation.)
372 Hand lay-up and bag molding vacuum bag and an expendable vacuum bag. There are often difficulties in having the facilities to handle large reusable vacuum bags weighing several hundred pounds. A cost evaluation must include all aspects of the program. There are some very large aircraft components utilizing expendable vacuum bag materials very successfully. The more complex the composite structure becomes, the more effective the reusable vacuum bagging system is.
are combined when the product is made. The composite designer must consider how the load bearing fibers are placed and ensure that they stay in the proper position during the fabrication. 16.6.1 DESIGN PROCESS
With some large composite structures, potential problem areas can best be identified using scale models. Working problem areas on an individual basis, a major factor in the success 16.6 DESIGN of a program is the amount of planning that The fundamental information needed for any can take advantage of the work force experidesign includes the stresses applied under ence. Successful composite structures are not storage and use and the strength of the mater- fabricated by one person; they require team ial used. Assume that the size, shape, quantity work from all disciplines. It is considerably and rate of production have dictated the use of harder to make a smaller composite structure, open mold techniques. Then the final thick- than make the full scale article. Procedures ness, orientation and quantity of reinforcing developed for the scaled article can, however, fibers are dependent upon the stresses that be easily translated into a full scale structure. must be resisted, how often and for how long. It is essential that the designer find out 16.6.2 FIRST ARTICLE FABRICATION what strength can be built into the laminate. This sets composite structures apart from The first article, also known as tool proof artiother types of materials, since the material is cle, can be used to provide information not made during fabrication of the product; the only on the tool to manufacture to the correct percentages and orientation of the reinforce- tolerances but also to produce an acceptable ment and the types of resins determine the composite structure meeting the design requirements. In addition, it can be utilized as properties of the final laminate. During the initial phase of the development a proof of the documentation of quality conof a composite structure, there is a need for trol inspection requirements, manufacturing design, manufacturing engineering, tooling, procedure's verification and allow design materials and quality control to provide inputs engineering to review overall requirements. so that the selection processes can be established. By coordinating early in a program, 16.6.3 DESIGN DETAILS one can focus on the real problems of design and manufacturing. The preplanning phase Parts with severe contour and thickness will allow for a program to develop at a more variations rapid pace. Confirming materials, manufacturing methods, tooling concepts and design It may seem easy to incorporate variations in requirements early can avoid the extra time contour and thickness into the design of a new and expense to attempt to make the composite product but in open mold products such variations must be made with caution. The fabrication process successful. The design of composite structures, while molding operation requires laying the materessentially similar to conventional design, ial on the mold to follow mold contour. If the does have the added dimension that materials angles are sharp (90' without radius), the lay-up
Design 373 will not follow the mold surface and will develop voids and resin rich areas in the laminate in the vicinity of the angle. For instance, in inside right angle corners without radius, (Fig. 16.12(a))the laminate will not pack into the corner. When there are sharp outside corners, the laminate (Fig. 16.12(b))will not wrap tightly over the corner. The solution to such problems is to design with a generous radius, preferably 4.75-12.75 mm (0.187-0.500 in) inside and out. The laminate will then follow the contour. Abrupt changes in direction are high stress areas and tend to delaminate and crack. They should be avoided and moderate self reinforcing curvatures used.
Changes in thickness To change thickness in open mold construction is to add or remove plies of material. An abrupt change means that the plies must be carefully laid up in a precise pattern. An abrupt change in thickness (Fig. 16.13(a)) results in a stress concentration and should be avoided as delamination is sure to occur at such a point. The solution to this problem is not to have abrupt changes but to gradually change by stepping back or 'shingling' the layup (Fig. 16.13(b)).
Openings The best opening is a round hole; the worst, an opening with sharp, non-rounded corners. The solution to stresses in an opening is to use large radii in the corners, to build up thickness gradually at the sharp corners, or to design a molded in flange around the opening (Fig. 16.13(c)).
1
L Fig. 16.12 Corner lay-up techniques (a) radius corner; (b) no radius corner.
I
Fig. 16.13 Changes in ply lay-up (a) abrupt changes; (b) stepped piles; (c) hole reinforcement.
374 Hand lay-up and bag molding
Joints and bonding Although common practice is to use joints intended for other methods of fastening, structural adhesives require joints of a special design. It cannot be stressed too strongly that the practice of using ordinary joints that have been slightly altered can lead to disastrous results. The type of joint used depends on the basic characteristics of the adhesive since structural bonds act over an entire area and not at a single point as rivet fasteners do. A joint should therefore be designed to minimize concentrations of stress. There are four basic types of stress encountered in structural bonding: tensile, shear, cleavage and peel. 16.6.4 DESIGN CONSIDERATIONS
However, these structures are not generally highly stressed structures. Large amounts of waste can be expensive for any program. As a rule, in programs where is little or no preplanning, the waste factor can be 20-35%. This may be acceptable for low cost materials such as fiberglass but for carbon prepreg materials costing over US$59OO/kg for 827 GPa (120 msi) pitch fiber, the cost for the waste can break an otherwise successful composite program. A well-planned approach to the cutting, kitting, lay-up and inspection requirements can reduce the waste factor to 10-15%. Large structures tend to have less waste than small components. The carbon epoxy central cylinder for a modern communication satellite has less than 3% material waste, the majority being for localized reinforcements. The America's Cup racing yacht was fabricated with less than 2% material waste.
The initial lay-up starts with the preparation of the prepreg ply kits. Individual ply kits will reduce the overall labor requirements during the ensuing lay-up. Hybridization of materials can be achieved on a ply-by-ply basis. Off axis 16.6.5 GENERAL DESIGN PRACTICE - . . - . (* 45") ply orientations can be prepared utilize Attention to ply orientation on strength ing a woven graphite cloth. Interleaf layers of controlled laminates can prevent matrix titanium or fiberglass can distribute high load and stiffness degradation. The 0" ply orieninputs into the laminate. tation is used to carry the longitudinal With large composite structures, i.e. racing loading, the 90" ply orientation the transyachts, the prepreg materials are usually disverse loading and the 45" ply for shear pensed directly onto the lay-up. One must loading. allow for 'fresh' prepreg, with maximum tack, e In order to minimize in-plane shear, place to be applied to the laminate with minimum the +45" and -45" plies together; the inout time of the material during the fabrication plane shear is carried by the tension and phase. compression in the 45" plies. One of the first reactions to using hand layTo minimize warpage and interlaminar up fabrication methods for a composite shear within a laminate, maintain the symstructure is: "this will be too labor intensive". metry about the center line of the laminate. This can be misconstrued to mean that all Stress concentrations can be minimized by composite structures done by the hand lay-up designing tapered or stepped laminate are expensive. Complex, integral, stiffened thickness changes. composite structures may not only be cost The placement of specific ply orientations effective but may not be able to be fabricated can influence the buckling strength and by other methods of manufacture. damage tolerance. The outer ply orientaCompression molding, resin transfer moldtions influence the laminate bending ing and sheet molding processes are cost characteristics more than plies placed at the effective for specific types of material forms. neutral axis.
-
*
Applications 16.7 APPLICATIONS
16.7.1 AEROSPACE
Two aerospace examples that were manufactured utilizing the hand lay-up and vacuum bagging process were an ICBM equipment section structure (300 parts) and a central cylinder (6 parts) for a modern communication satellite. These two primary composite structures illustrate that hand lay-up and vacuum bag procedures can be effectively utilized on limited production programs. For greater production demand or with a less complex structure other methods (machine lay-up or filament winding) could have been used. However, neither alternative process could achieve the results required with materials selected.
ICBM equipment structure The ICBM equipment structure was initially designed to utilize unidirectional graphite/epoxy materials because unidirectional materials were available during the design phase. The development of woven graphite was undertaken in order to reduce the manufacturing costs of the program. Initial prototypes (unidirectional tape) required over 2000 man-hours labor to complete. The first prototype utilizing woven graphite required only 900 man-hours, better than 50% reduction in manufacturing costs. Eventually production labor requirements were reduced to less than 200 man-hours. Satellite central cylinder The graphite/epoxy central cylinder was a unique structure from its inception. The project was undertaken not just because composites would be lighter but that the metallic (beryllium) design required a longer manufacturing time that was unacceptable. The materials of choice were a 520 GPa (75 msi) pitch graphite fiber, epoxy resin and aluminum honeycomb. During the initial
375
manufacturing/design reviews, the decision to fabricate the central cylinder as a one piece monococque structure was made. One quarter scale models were made as initial feasibility trials and the finished cylinder was 108 cm (42 in) in diameter and 231 cm (91 in) high. With that initial success, a full size manufacturing development cylinder required 1700 man-hours to fabricate. By production unit five, the fabrication time was reduced to 500 man-hours per cylinder. The final design incorporated compression molded graphite inserts for hard mounts and was co-cured as a one piece structure. The program set new standards of cooperation between engineering analysis/design, manufacturing and quality control groups. A manufacturing plan including all required inspection points and a detailed fabrication sequence was prepared. The major problem during the fabrication was the out time available with the graphite epoxy resin system. With only 10 working days available, the kit preparation, lay-up, compacting cycles and final cure was on a tight time schedule. This program used various operational procedures to achieve success. Tank inserts were compression molded from chopped carbon/epoxy material prior to the start of the laminate schedule on the cylinder. It was determined early in the development cycle that the unidirectional tape material handled better if it was not precut into kits, but prepared just prior to application to the tool. A combination of heat debulks and pre-bleeding was utilized to maintain the desired resin content, until the supplier could prepare an acceptable net resin unidirectional tape. With limited facilities to autoclave a large cylinder with in-house capability, other methods of manufacture were employed. After the completed hand lay-up was done, a vacuum debulk was applied to ensure all air was removed and the lay-up was compacted. The cylinder was then wrapped with perforated shrink tape (to allow for resin/air bleed during cure), then breather cloth was applied
376 Hand lay-up and bag molding Challenge. A typical America’s Cup yacht utilized over 13 006 m2 (140 000 ft’) of ply surface area of unidirectional graphite tape in the hull and deck structure and 9290 m’ (100 000 ftz)of ply surface area for a one piece graphite mast structure. Due to the overall size and past experience with the boat builder, a male wood plug was fabricated with integral heated wires imbed16.7.2 MARINE APPLICATIONS ded to reduce the heat sink effect during the The fabrication of an America’s Cup racing final co-cure of a complete hull laminate. As yacht (Fig. 16.14) presented another set of the maximum temperature allowed was 90°C requirements. With limited cure temperature, (183”F),wood was a good choice. The hull laminate was then applied, startresin matrix and control density requirements, ing with the inner skin (as on an America’s the challenge was to be able to hand lay-up Cup racing yacht) then film adhesive and honand cure a large composite structure to meet eycomb core was applied. The inner laminate design requirements. Marine applications had reached new heights when the new rules went and honeycomb was vacuum bagged and parinto effect for the 1992 America’s Cup tially cured. The honeycomb core was smoothed and all joints were filled prior to the application of the outer skin laminate. The outer laminate and any local reinforcements were applied to the core and inner laminate (Fig. 16.15). The completed hull laminate was then vacuum bagged and oven cured. The outside of the hull was essentially complete but had a rough surface. The roughness was greatly dependent upon the care of workers during the lay-up of the outer laminate skin. With reasonable precautions, prepreg material can be placed in such a way that there are no overlaps and all gaps have been filled with additional fiber. This effort alone can save hundreds of man-hours during the final finishing. This effort saved over 1200 man-hours compared with a ’wet’ prepreg that did not produce an acceptable outer surface finish. The oven cure achieved maximum mechanical properties required by the design. Laminate
prior to final nylon vacuum bag. Studies were conducted on utilizing a silicone rubber vacuum bag, but overall program costs and difficulties in handling a large (heavy) bag, pushed the utilization of nylon bagging film for the final vacuum bag. The part was then oven cured.
Fig. 16.14 ’Spirit of Australia’.
consisted of filling and grinding the surface in
Applications 377
-
1-
----
-
.k& I
I
*
I
t Fig. 16.15 'Spirit of Australia' laminate laydown (top) 45" ply laminate; (bottom) 0" ply laminate.
two or three stages and culminating with a final coat of good grade epoxy or urethane marine paint. This step is often postponed until the boat is completely assembled and ready for fitting out. Then, the hull was inverted and placed in a fitted saddle; the male plug was removed. Additions of bulkheads, flooring and interior completed the work. The manufacture of the carbon/epoxy mast for the 'Spirit of Australia' presented some unique challenges. The major program challenge was how to obtain a mast that would provide the required performance at affordable
cost. A discussion with design/analysis and manufacturing concluded that it was possible to fabricate a one piece mast within the facilities available and technical expertise within the syndicate. The design of the tooling was aimed at providing a capability to cure the carbon/epoxy resin within the America's Cup rules. The requirements specified the cure temperature 120°C (250"F), cure pressure (3 atmospheres), laminate density /modulus of graphite fiber and overall mast profile. All these objectives were realized within 100 days, from start to completion of the first one piece, 35 m (115 ft) in length, America's Cup composite mast. Performance exceeded all expectations. Additional masts starting with mast number 3 were manufactured with the final weight objective of less than 450 kg (990 lb). Total fabrication time for each mast, from the start of material kit preparation to completion of the cure, required a maximum of 21 days. The selected carbon/epoxy resin system was workable for this limit. The mast, spinnaker pole and rudder stock utilized variations of pressure bag molding with integrally heated fiberglass tooling. These two tooling approaches allowed for some of the largest one-piece structures to date to be fabricated in the commercial marine market. Logistics can become a major role player in the planning required in that not only the support materials are on hand but that power is available; materials handling is taken into account and there is sufficient crew available to complete the lay-up within the material out time limits. As seen in the above examples, the one major factor for all the programs was the material out time. With the development of new resin matrixes, the design of complex structures can be achieved.
MATCHED METAL COMPRESSION MOLDING OF POLYMER COMPOSITES
17
Enarnul Haque and Burr (Bud) L. Leach
17.1 INTRODUCTION
In today’s highly competitive global economy, the need for materials with the right properties to meet the demands of design, environment, durability and economics is growing. Composite materials, with their high strength and stiffness-to-weight ratios, have many advantages and are a desirable engineering material. There is no universal definition of composites. In general, a composite material is a heterogeneous material system consisting of two or more physically distinct materials. In a composite material system, the individual materials exhibit their unique properties and the composite as a whole shows properties that are different from its constituents. In addition to the constituents’ unique properties, the properties of composites are also dependent on the form and structural arrangements of the constituents and the interaction between the constituents. Broadly speaking, composites consist of two components, a binder or matrix and a reinforcement. The matrix functions as the body constituent, serving to bind the reinforcement together and giving the composite its bulk form. The reinforcements are the structural constituents, providing high strength to the internal structure of the composite.
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Reinforcement may take the form of fibers, particles, laminate, flakes and fillers. Depending upon the type and orientation of the reinforcement and the manufacturing technology required to produce them, composites with various properties and cost can be fabricated. Polymer composites are composites in which the binder or matrix is a polymeric material and the reinforcement is usually a thin fibrous material. Polymer composites can have either a thermosetting or thermoplastic matrix. In this chapter we will discuss thermosetting matrix based composites. Reinforcing fibers may also be of various kinds, with glass (E-type), carbon, or organic fibers (e.g. aramid) being the most common. Glass fibers are the most widely used type of reinforcement since they offer good strength and moderately high temperature resistance (about 260°C) at a cost effective price. Glass fibers also come in various forms. They can be continuous filaments, cut or chopped strands, roving and yarns, or in the form of cloth, mats or tapes. Thus allows glass fibers to be used in a variety of applications such as lay-up, filament winding, matched die-molding, etc. In this chapter we discuss the matched metal compression molding of thermoset based polymer composites. Matched metal compression molding is a molding process in whch the cure is obtained while the material is restricted between two mold surfaces and the loading and closing of the mold causes the material to conform to the desired configuration. This
Background 379 process enables large scale production of large surface area parts with contour problems and tight tolerances. Matched metal compression molding employs a ’mold’ or match dies. The male mold is matched to the female mold so that when the dies are closed, a controlled space results. A preform charge is placed on the core and the cavity is pressed against it, applying direct pressure on the material. The pressure in this type of molding varies from 1.38 to 6.895 MPa (200 to 1000 psi) and curing temperatures from 125°C to 160°C (260°Fto 320°F).
17.2.1 BULK MOLDING COMPOUND (BMC)
BMC has been defined as ‘a fiber reinforced thermoset molding compound not requiring advancement of cure, drying of volatile, or other processing after mixing to make it ready for use at the molding press’*. BMC can be molded without reaction byproducts under only enough pressure to flow and compact the material. BMC is usually manufactured by combining all the ingredients in an intensive mixing process. Recent advances in BMC technology dictate that both the dry ingredients and wet ingredients be batch mixed separately and then combined 17.2 BACKGROUND together in an intensive mixer. The BMC is usuAdvanced polymer composites are now being ally in a fibrous putty form when it comes out of applied extensively for all types of applications the mixer and resembles ’sauerkraut’. It is usuin the industrial and automotive markets. Table ally compacted and extruded into bars or ’logs’ 17.1’ shows the usage of composites in various of simple cross section. The earliest BMCs were probably made markets during 1991-1993. This section deals primarily with thermoset polymer composites about 1950, employing a process of impregnatused in matched metal compression molding. ing roving strands with blend of resin, filler, etc. The two most popular reinforced molding com- and chopping them to a length in the wet stage. pounds used in the plastics industry are Premix Since wetting glass fibers with a resin containor BMC and SMC (also referred to in modified ing much filler is difficult and slow, these versions as HMC and XMC). Low Pressure premixes had a high glass content. The first Molding Compounds (LPMC),ZMC and TMC high volume commercial BMC was made with sisal fibers and used in molding automobile are also becoming popular.
Table 17.1 US Composites shipments: 1991-1993”
Millions of pounds Markets Aircraft/ aerospace /military Appliance/business equipment Construction Consumer products Corrosion-resistant equipment Electrical/electronic Marine Transportation Other Total
1991 38.7 135.2 420.0 148.7 355.0 231.1 275.0 682.2 73.8 2359.7
1992 32.3 143.2 483.0 162.2 332.3 260.0 304.4 750.0 83.4 2550.8
1991-1 992
1993
1992-1 993
% change
(projected)
% change
-16.5 +5.9 +15.0 +9.1 -6.4 +12.5 +10.7 +9.9 +13.0 +8.1
26.0 146.7 522.0 164.1 336.8 273.0 317.2 810.0 88.0 2683.8
-1.95 +2.4 +8.l +1.2 +1.4 +5.0 +4.2 +8.0 +5.5 +5.2
a Includes reinforced thermoset and thermoplastic resin composites, reinforcements and fillers. Source: SPI Composites Institute
380 Matched metal compression molding of polymer composites heater housings. Improvement in the binder chemistry of glass fibers, development of a chemical thickening system and thermoplastic low profile additives help BMC to attain strength, chemical resistance and to overcome surface irregularities. Consequently, BMC was accepted for use in the electrical, chemical and appliance industries. Today, BMCs are accepted as high performance engineering thermoset molding compounds and used extensively in the electrical, automotive and consumer goods industries. BMC is increasingly injection molded to take advantage of the automation and reproducibility afforded by the process, although it is also both transfer molded and compression molded.
17.2.3 THICK MOLDING COMPOUND (TMC)
TMC was developed by Takeda Chemical Industries, Ltd (Osaka, Japan). TMC is suited to compression, injection and transfer molding and is usually processed on the same equipment as SMC and BMC materials. TMC composites are usually produced up to 51 mm (2in) thick and glass fiber length can vary from 6.4 to 50.8 mm (0.25 to 2 in). In TMC, continuous impregnation and high sheet weight result in complete wet-out of resins, fillers and reinforcing fibers. Better wet-out results in improved mechanical properties and reduced porosity. TMC is usually used in business machine housings, appliance components and other consumer related industries.
17.2.2 Z MOLDING COMPOUND (ZMC)
17.2.4 SHEET MOLDING COMPOUND (SMC)
ZMC was developed in 1979 in France to improve BMC performance. BMC suffers from glass fiber degradation during injection molding and ZMC was developed to keep shear forces as low as possible during moldin$. A special type of injection molding machine developed by Billion in France combines the advantages of both a screw machine and a plunger machine. The ZMC injection machine uses a screw to homogenize and precisely measure the shot. The injection is made like a plunger by the displacement of the screw and inner barrel inside the main barrel. In a ZMC, the different components are mixed in conventional mixers like BMC. The compound viscosity is usually low and adapted to injection machine characteristics. The design of the mold plays a key role in the processing phase and ZMC parts cannot be successfully made unless part design and mold design are combined upfront. Compared to SMC, ZMC parts have lower mechanical properties, but higher performance when compared to conventional injection molded BMC.
SMC is a type of fiber reinforced plastic which primarily consists of a thermosetting resin, glass fiber reinforcement and filler. Additional ingredients such as low-profile additives, cure initiators, thickeners and mold release agents are used to enhance the performance or processing of the material4. The development of SMC started in the early 1950s after the finding that the viscosity of unsaturated polyester resins increases with the addition of Group IIA metallic oxides, hydroxides, or carbonates5. The first published report on SMC was presented at the Cleveland Section of the Society of Plastics Engineers meeting. The report involved work done in Germany using fiberglass mat impregnated with a resin mixture containing magnesium oxide6.At the same time a number of US patent^^,^ were published on the use of Group I1 metal oxides, hydroxides, or carbonates for use on adhesives. The early applications of SMC materials were in electrical and industrial goods. During the next two decades, growth in commercial usage of SMC followed the evolution of continuously improving equipment, low profile additive,
Background catalyst, etc. The automobile industry started using SMC in the early 1970s for producing exterior body components, such as hoods or grille opening panels. With the introduction of high strength SMCs in the mid-l970s, usage of SMC increased to structural components. SMC is currently used extensively in transportation, construction (door panels), appliances (washing machine door, refrigerator housing), furniture (chair, tabletop) and business machines (computer housings). The transportation industry has the highest level of consumption of SMC. For instance, in the North American market alone, the annual rate of consumption exceeds 100 million kg9. Details of SMC manufacturing are available in the literature5.SMC offers many advantages which include variety, part consolidation, lightweight and dimensional stability. With the evolution of flexible backbone polyester resin systems and development of special additives, flexible SMC is becoming very popular and is now competing with thermoplastics for vertical body applications. Special applications SMC is also becoming popular. With the addition of hollow microsphere glass bubbles in a standard SMC formulation, lower density (1.3-1.4) is obtained for weight reduction. High strength molding compound (HMC) is a SMC containing 65% chopped glass fiber instead of the usual 25-35%. HMC uses little or no filler and can be compounded on a standard SMC machine. Directionally reinforced molding compound (XMC) is a directionally oriented moldable resin-glass fiber sheet containing 65-75% continuous reinforcement. XMC is also usually compounded on standard filament winding equipment and has strength five times greater than SMC. Unidirectional molding compound (UMC) is a system of chopped and continuous fibers produced on a modified SMC machine. An advantage of UMC is that different varieties of fibers can be used.
381
17.2.5 LOW PRESSURE MOLDING COMPOUND (LPMC)
Low Pressure Molding Compound (LPMC) is an SMC type material which can be molded at 1.38-2.07 MPa (200 to 300 psi) instead of 5.52 -6.90 MPa (800-1000 psi) required for standard SMC. LPMC is made by replacing the chemical thickening mechanism of alkaline earth oxides (Group 11) with a physical thickening mechanism utilizing a crystic polyester. The material is heated to melt the crystic and then the other ingredients are added, mixed together and run through a modified SMC machine maintaining the elevated temperature. The thickening occurs as the material cools to ambient temperature and the compound is ready to mold at that time. Cooling rolls speed the cooling process and thus the material can be molded right off the SMC machine without waiting for the 2 4 4 8 h thickening of standard SMC. LPMC allows the molder to use lower tonnage presses to mold larger parts and use less steel in building the tools as they do not have to deal with high pressures and corresponding forces. The shelf life of LPMC is much longer than SMC and the physical properties are comparable. 17.2.6 CONTINUOUS IMPREGNATED COMPOUND (CIC)
In 1986, continuous impregnated compound (CIC) was developed in Germany. This is similar to TMC. Like TMC, the impregnation is made between two rolls but the compound is removed by doctor blades and carried by a screw or plunger to boxes or drums. CIC is usually injection molded, but can also be injection/compression molded. Properties are comparable to BMC but processing is easier than BMC. Modified CIC is also known as KMC (Kneaded Molding Compound).
382 Matched metal compression molding of polymer composites 17.3 FORMULATIONS
Polymer composites have the unique and distinct advantage in that their properties can be tailored to meet different applications by designing the formulations. The major components of polymer composites used in matched metal compression molding are resin, low profile additive, fiber, filler, initiator, inhibitor, internal mold release agent and other additives (e.g. viscosity reducer, toughness enhancer, etc.).
butadiene copolymers. Extensive details of LPA mechanism are published in the literature9J8J9. 17.3.3 INITIATORS Ah-D INHIBITORS
Initiators are used to initiate the curing reaction at elevated temperatures. Composites are polymerized or crosslinked by a free radical mechanism in which the double bond of the polyester chain reacts with the vinyl monomer (usually styrene). This copolymerization reaction provides a three-dimensional network that converts the viscous liquid resin 17.3.1 RESIN paste into a hard thermoset solid. Initiators Unsaturated polyesters and vinyl esters are the decompose at elevated temperature and provide a source of free radicals to initiate the principal resins used in polymer composites for compression molding. Epoxies are also used for copolymerization reaction. Peroxyesters and specialty products which require longer cure peroxyketals are the most common classes of cycles and higher strength. Phenolics are being peroxides. Inhibitors are added in small quantities to used for formulating composites, especially prolong shelf life, modify cure rate and magSMC, in applicationswluch require lower flamnitude of exotherm to prevent cracking of mability, reduced smoke generation and higher thick molded sections. Inhibitors are also thermal stability’O. Details of resin chemistry used to improve resin stability. Two general are available in the literaturel’-l4. Styrene is classes of inhibitors are commonly used, subcommonly used for cross-linking of both polystituted phenolic derivatives and the ester and vinyl ester resins. Low styrene quaternary ammonium salts (e.g. hydropolye~ter’~ is becoming popular due to strinquinone, p-benzoquinone, etc). An excellent gent EPA requirements on styrene vapors. New review on initiator and inhibitor chemistry is resin technology is also being considered for available elsewhere y. compression molding. They include hybrids of unsaturated polyester and urethane16,acrylesterol resin with polyi~ocyanate’~, etc. 17.3.4 FILLERS
Fillers are used to improve physical properties, reduce volumetric shrinkage of the resin and to reduce costs. Fillers are typically divided into Low profile additives are thermoplastics that functional and non-functional categories. are added to the formulation in 2-5% (by Examples of functional fillers include alumina weight) of the final product or 10-20% (by trihydrate for flame retardancy, hollow glass weight) of the organic portion of the formulabubbles for lower weights, mica and wollastion to control the shrinkage of the cured tonite for reinforcement. Non-functional fillers composites. Typical thermoplastics include are used for cost reduction and are mineral polyvinyl acetates, poly methyl methacrylate based. Ground limestone (CaCO,) is the most and copolymers with other acrylate, vinyl chloAn excellent review is common type of filler. ride-vinyl acetate copolymers, polyurethane, available on filler use in composite^^^^^^ polystyrene, polycaprolactone, cellulose acetate butyrate, saturated polyester and styrene17.3.2 LOW PROFILE ADDITIVES
Molding 383 17.3.5 FIBERS
17.3.8 OTHER ADDITIVES
Glass fiber reinforcement is used to achieve Pigments are added to produce color in the necessary dimensional stability and mechani- molded part. Common pigments include cadcal properties. E-glass is the most common mium salts, carbon black, titanium dioxide, fiber reinforcement for composites. iron oxides, organic dyes and pigments, etc. Depending on the binder chemistry and Various types of viscosity reducers are used to amount, glass fibers are classified as hard or lower the viscosity of the paste to increase soft type. Other types of fibers include carbon, filler loading and glass wet-out. Other addiaramid (Kevlar), S-2 glass, etc. Glass loading tives include various elastomeric additives normally averages 30% by weight in compres- (e.g. Hycar, Kraton, etc.) to increase toughness. sion molded composites, but can vary from Tables 17.3 to 17.6 show typical formula18-65?” by weight. Table 17.2” shows typical tions for BMC, SMC, ZMC and LPMC. fibers used in polymer composites. 17.4 MOLDING
17.3.6 INTERNAL MOLD RELEASE
Internal mold release agents are added to facilitate part ejection from the mold. Major types of release agents include metallic stearates, fatty acids, fatty acid amides and esters and hydrocarbon waxes. Zinc stearate and calcium stearate are the most widely used internal mold release agents in SMC and BMC. 17.3.7 THICKENERS
The addition of Group I1 oxides and hydroxides to carboxy-terminated unsaturated polyester/vinyl ester resin increases its viscosity. Magnesium oxide, magnesium hydroxide, calcium oxide, calcium hydroxide or combinations of those materials are the most popular thickeners.
Matched metal compression molding is one of the oldest manufacturing techniques in the plastics/composites industry. The recent development of high strength, fast cure, Table 17.3 BMC formulation
PHR Polyester resin Low profile additive Styrene Initiator Inhibitor Mold release Pigment Thickening agent Filler Compound Glass fiber Paste
60.0 40.0 5.0 1.5 Trace amount 4.0 0.25 1.0 50-200 10-25% 75-90%
Table 17.2 Glass fibers used in compositesz1 Fiber
E-glass S2-glass Carbon (graphite) Kevlar 49
Specific gravity
Tensile strength (GPa)
Tensile modulus (GPa)
Tensile failure strain (%)
Coeficient of thermal expansion (x 1 @6/OC)
2.54 2.48 1.76-2.15
3.45 4.30 1.5-5.6
72.4 86.9 220-690
4.8 5.0 0.3-1.2
1.45
3.62
131
2.8
5 2.9 -0.1to -1.2 (longitudinal) 7-12 (radial) -2 (longitudinal) 59 (radial)
384 Matched metal compression molding of polymer composites Table 17.4 SMC formulation
PHR Polyester resin Low profile additive Styrene Initiator Inhibitor Mold release Pigment Thickening agent Filler Compound Glass fiber (25.4 mm) Paste
55.0 40.0 5.0 1.5 250 ppm 4.0 1.0 2.0 150-250 25-30% 70-75%
physical and mechanical properties can be obtained in compression molded parts. Figure 17.1 shows a schematic of a compression molding process. This section addresses the compression molding of composite parts using SMC. BMC molding is similar except for the charge preparation step. The compression molding process can be divided into four distinct steps. Heat and pressure
+
Cavity
Table 17.5 LPMC formulation
PHR Polyester resin Crystic Styrene Initiator Mold release Pigment Filler Compound Glass fiber Paste
65.0 15.0 5.0 1.2 5.0 1.2 220 4
25-30% 70-75%
Table 17.6 ZMC formulation
PHR Polyester resin Low profile additive Styrene Initiator Inhibitor Mold release Filler Compound Glass fiber Paste
65.0 40.0 5.0 1.5 100 ppm 4.0 220
SMC/BMC and advancement in press technology is making the compression molding process very popular for mass production of composite parts. In comparison with the injection molding process, in general, better
I Heat and pressure
Fig. 17.1 Schematic of a compression molding process.
17.4.1 CHARGE PREPARATION AND PLACEMENT
When the SMC has reached its desired molding viscosity, pieces of SMC are cut to pre-specified size after removing the carrier films. The SMC is cut using slitters, pizza type or guillotine type cutters. Several pieces of SMC plied together form the 'charge'. The charge pattern/ply dimensions are chosen so as to cover 20-80% of the mold surface area. The charge pattern and placement on the mold determines the quality of the molded parts, since it influences the length of flow in the mold, fiber orientation,
Properties 385 flowline and other surface defects. In order to reduce cycle time, sometimes the charge is preheated to a temperature below gel point using infra-red or dielectric heaters. 17.4.2 MOLD CLOSING AND FILLING
entire surface. After IMC injection, the press is closed and the curing operation is repeated at or above the SMC molding pressure. Sometimes the IMC is injected at high pressure without mold opening and closing prior to complete cure of the SMC charge.
After proper placement of the charge in the core of the mold, the cavity is quickly closed to 17.4.4 PART EJECTION AND POST-CURE contact the top surface of the charge. The cavAt the end of the molding cycle, the mold is ity is then closed at a slower rate, usually 4-12 opened and the part is ejected from the core mm/s. In most cases the mold is heated to (for with the use of integral ejector pins and example) 150°C, which causes the charge visallowed to cool to ambient temperature. Hot cosity to be reduced. With increasing mold parts are handled carefully and are usually pressure as the mold is closed, the charge placed on a support racks to cool to ambient flows towards the cavity extremities, forcing temperature. As the part cools outside the air out of the cavity. The mold closing speed is mold, it continues to cure and shrink which very important as it induces gelation of the top creates residual stresses due to differential charge surface if the closing speed is slow or it cooling at various sections in the part. After causes trapped air if closing speed is fast. The the part is placed on support rack, it is filling stage is usually completed in 0.5-20 s5. deflashed while still hot and stored in racks for Vacuum molding is increasingly being used secondary operations like punching, drilling, during charge flow to reduce surface porosity bonding, etc. and air entrapment in the part. Vacuum level The compression molding process is comis usually in the range of 7-9 x lo4Pa (21-27 in plex and there are several important variables Hg). The molding pressure based on projected that influence molding. Compression molding part area ranges from 1 to 10MPa (100 to may also produce a variety of surface and 1200 psi). Higher molding pressure causes internal defects which can be eliminated by sink marks, while lower pressure cause scumproper material selection, part design and ming of the mold and porosity. molding technique. Details of the molding variables and the source and remedies of 17.4.3 CURING major molding defects are available in the literature5,'. After filling, the charge remains in the hot mold for the crosslinking reaction to be completed. The curing time is usually between 25 s 17.5 PROPERTIES to 3 min, but depends on several factors, including resin-initiator-inhibitor reactivity, The properties of a polymer composite can be tailored, within limitations, to meet different part thickness and mold temperature. Sometimes in class A or appearance grade applications by designing its formulation. This parts, in-mold coating (IMC) is used to unique characteristic of polymer composites enhance the surface of a molded part. The makes definition of detailed properties diffimost common method of IMC injection cult. The properties are usually used for requires opening the mold by a small amount information and guidelines for preliminary (0.2-0.5 mm) after the curing cycle. IMC is part design, material selection and to underusually a coating of polyester or stand the effect of formulation variables on polyester-urethane hybrid which covers the mechanical properties.
386 Matched metal compression molding of polymer composites 17.5.1 STATIC PROPERTIES
17.5.3 OTHER PROPERTIES
Table 17.7 shows the static and impact properties of SMC, BMC, ZMC and LPMC. In general, tensile and flexural properties are routinely measured and are presented here. Compressive and shear properties are measured only for use in special applications. The static properties of SMC and BMC are highly dependent on the fiber content, length, type and orientation. Tensile strength increases significantly with increasing fiber content; however, the tensile modulus is affected only moderately. Increasing the length of chopped fiber increases the tensile strength, but has no effect on the modulus. Glass fiber type (E-glass or S-glass) has a significant effect on both the tensile strength and modulus. The resin chemistry also influences tensile properties at low fiber content. In general, flexural and compressive properties follow the same trend as the tensile properties. Flexural strength is always higher than tensile strength, though the modulus may be comparable.
Several other tests are now being performed to correlate properties with operation conditions. The dynamic mechanical analyzer (DMA) is used to measure complex and storage modulus at various temperatures and frequency ranges. The effect of environmental conditions on various properties is tested to simulate end-use environment. Creep and stress relaxation tests are also done on SMC/BMC for use in structural applications. Electrical properties are also important, permitting BMC to be used in electrical applications. Arc resistance is important and dielectric strength, dielectric constant, dissipation factor, etc. are also measured.
17.5.2 FATIGUE PROPERTIES
The fatigue properties of SMC and BMC are usually based on tensile cyclic loading of unnotched specimen. A typical S-N diagram is shown in Fig. 17.221.In general, the fatigue strength increases with increasing fiber content and there is no fatigue limit, unlike low carbon steel. Details of such testing are published elsewherez1,".
17.6 APPLICATIONS
Reinforced composites materials offer the maximum design versatility and capability of any material. With the excellent cost/performance characteristics of reinforced composites, the variety and quantity of products being produced with these materials grow annually around the world. Matched metal molded reinforced composites should be considered when the finished product can be enhanced by one or more of the following characteristics. Part consolidation Reinforced composites can be molded in three dimensions in one operation. Complex shapes that require multi-piece assembly using materials, such as wood or steel, may be molded in one step with the use of ribs, bosses and varying wall thickness.
Table 17.7 Static and impact properties"
Property
Tensile strength (MPa)
Tensile modulus (GPa)
Flexural strength (MPa)
Flexural modulus (GPa)
IZOD impact (unnotched)
Specific gravity
Coeficient of thermal expansion (x 1 P P C )
SMC BMC ZMC LPMC a
65-100 30-70 30-70 65-100
Published industry data
9.5-14 8-12 8.5-12.5 9.5-14
130-200 50-150 50-150 120-200
8-14 9-1 7 7-12 8-14
600-1200 100-700 200-500 600-1200
1.3-2.0 1.7-2.1 1.8-2.0 1.8-2.0
8-14 15-20 11-27 7-10
Applications 387 100 -
R = 0.05 80 -
a" z 6 v)
-
60-
cn
E 2 .-E
40-
X
r" 20 -
01 0.1
I
I
I
I
I
I
I
1
10
102
103
104
105
106
Number of cycles, N Fig. 17.2 Typical fatigue !+N diagram for SMC (21) A: at 40°C; 0:at 23°C; Cl:at 93°C. [Reproduced from Composite Materials Technology: Processes and Properties (ed P.K. Mallick and 5. Newman) by permission of
the publisher.]
Light weight Reinforced composites offer a greater strength-to-weight ratio than most non-reinforced plastics and many metals. Dimensional stability Reinforced composites can maintain the critical tolerances required of the most demanding applications. Composites meet the most stringent material stiffness, dimensional tolerance, weight and cost criteria in many diverse applications. High strength Reinforced composites have excellent strength-to-weight properties. By weight, reinforced composites surpass the tensile strength of iron, carbon and stainless steels. Many glass reinforced compounds equal or exceed the flexural strength and impact resistance of metals23.
from most organic chemicals and can be formulated to resist acidic and basic solutions. Electrical resistance Reinforced composites are very poor conductors of electricity. As such, they have a lvgh dielectric strength for application in the electrical and electronic industriesz3. Resistance to minor impact Reinforced composite components have an excellent memory characteristic. Instead of yielding or deforming under minor impact as with steel, a reinforced composite panel will deflect and spring back to its original surface form4(Fig. 17.3).
Surface quality Reinforced composites can achieve a variety of surface textures, from very smooth and glossy to a rough texture. Corrosion resistance Reinforced composites Insignias and alphanumeric characters can be do not rust or corrode, are resistant to attack molded as raised or indented characters.
388 Matched metal compression molding of polymer composites and performance requirements of the product or component. The designer must:
Steel
-
///// Composite
Fig. 17.3 Minor impact.
Molded-in color Color can be added to the reinforced composite compound, often eliminating the need for a secondary painting process. Recycling Most reinforced composites can be recycled either by regrinding or pyrolysis. Reground material can be used as filler or reinforcing material. Pyrolysis reduces the composite into its basic components by heating the material in the absence of oxygen. The process yields gas, oil and solid by-products that can be recycled back into composites, or used in building and agriculture materials 4. Thousands of products are molded each year utilizing reinforced composites: aerospace, automotive parts, sports and recreational equipment, boats and business machines to name a few. This wide variety of applications is indicative of the versatility, capability and cost effectiveness of reinforced composites.
1. establish size and shape limitations based on: 0 basic end use function; 0 aesthetics and marketing; 0 shipping limitations; 0 weight requirements; 0 strength and stiffness requirements; 0 flexibility requirements or limits; 0 process limitations. 2. establish the structural requirements based on: various loads that will be impacted to the part including weight, pressure and dynamic loads; duration of the loads on the part; temperature variations on the part and surface; number of cycles of temperature change; liquid, moisture and vapor resistance requirements; relative significance of strength-toweight ratios.
17.7 DESIGN CONSIDERATIONS
3. establish the non-structural requirements based on: 0 corrosion, weathering, moisture and temperature resistance; 0 moisture and vapor penetration for condensation protection; 0 fire safety relative to combustibility; 0 flame-spread rate requirements; 0 light transmission (transparency,translucency and opaqueness); 0 surface textures, both aesthetic and functional; 0 surface coatings for protection or aesthetics; 0 thermal insulation; noise and sound control; 0 dielectric requirements for electrical insulation24.
Given the wide range of options provided by reinforced composites, it is imperative that the designer accurately establish the functional
With the establishment of the functional and performance requirements the product design can be developed.
Design considerations 389 There are some general design principles which can assist in the development of structurally efficient configurations for reinforced composite components. 17.7.1 SHELL AND PLATE CONSTRUCTION
These are the most common configurations of reinforced composite parts. A reinforced composite component is constructed from layers of reinforced composite materials molded into the shape desired creating a geometric 'shell'. It is good practice to design so that only forces that place a part in tension or compression are applied to any component. Compound curve shapes provide good transmission of uniform loads into tensile and compressive forces within a part. Ribbed configurations are often used to achieve required strength and stiffness in structural components (Fig. 17.4). Corrugated or open ribbed configurations are used to achieve needed structural depth while efficiently using materials and fabrication processes (Fig. 17.5). 17.7.2 DRAFT
Draft is a slight angle introduced relative to the direction of the opening and closing of the mold. It is necessary to design the part so that all side walls, both interior and exterior, have draft. This enables the part to be removed from the mold without hanging or rubbing, which
Fig. 17.4 Rib configuration.
Circular
jf
Pitch
)t '
Rectangular
Fig. 17.5 Corrugated configuration.
can degrade 'appearance' surfaces (Fig. 17.6). Minimum draft angle of 1" for the first 76.2 mm (3 in) of depth, 2" for 76.2-101.6 mm ( 3 4 in) of depth, 3" for 101.6-127 mm (4-6 in) of depth and 1" for every additional 50.8 mm (2in) thereafter is recommended on all surfaces parallel to the mold movement. This pertains to all part details, such as ribs, bosses, elevation changes and holes. A draft angle of 1" on standing ribs and bosses will yield a thickness change of 0.43 mm (0.017 in) per inch per sidez4. 'Zero draft' may be obtained by designing the mold in such a way so the draft-free surface lies at an angle to the mold direction. This will affect the positioning of bosses, ribs and other details of the part.
390 Matched metal compression molding of polymer composites Draft angle (1’ recommended) 1Ae” Minimum Recommended
‘By
Radius Determined Part Thickness
Fig. 17.8 Minimum draft. mold movement
Fiig. 17.6 Draft.
17.7.3 RADIUS
17.7.4 NOMINAL THICKNESS
In mold making, the radius defines the curvature established between two intersecting surfaces. The more generous the radius, the better the flow of molding material for a stronger part (Fig. 17.7). A minimum radius of 1.59 mm (1/16 in) is recommended for all radii for both interior and exterior plane intersections. Radii should be designed to maintain relatively uniform part thickness (Fig. 17.8). Ribs and bosses opposite an appearance surface should have the radii eliminated to reduce the likelihood of warpage or ’sink’ (surface depression).
The nominal thickness is the overall design thickness of most of the part. It is desirable to establish uniform thickness throughout a part, to achieve minimum cure time, uniform cooling and minimize warpage and shrinkage (Fig. 17.9). Nominal thickness for reinforced composites is 2.544.57 mm (0.100-0.180 in). Recommended minimum thickness is 1.53mm (0.060 in). Recommended maximum thickness 25.4 mm (1.00 in). By designing hollow ribs, bosses and elevation changes can achieve intricate part geometry while maintaining nominal thickness throughout the part. 17.7.5 EDGE STIFFENING
Fig. 17.7 Outside radii.
Edge stiffening is a design characteristic applied to unsupported edges to prevent warping or bowing. Edge turning is preferable to edge thickening due to the possibility of porosity at the edge of part caused by the lack of molding pressure on the thick area (Fig. 17.10).
Design considerations 391
Fig.17.9 Nominal thickness. Uniform thickness promotes uniform flow and curing and minimizes the risk of warpage, distortion and telegraphing at thickness changes through the surface. 17.7.6 RIBS
Linear projections 90" from the plane surface of a part are called ribs. The use of ribs will allow the part to meet strength and rigidity requirements, preventing warpage and bowing in large plane surfaces while reducing the bulk and mass of a part. Ribs should be designed to maintain the nominal thickness and follow the guidelines
for draft angles. They should be dimensioned so that their thickness at the juncture of the rib with its plane surface is between 75 and 90% of nominal (Figs 17.11 and 17.12).
4-
0.5"draft
Fig. 17.11 Rib geometry for class 'A' surfaces.
(0.06'') radius
+1.0" draft Fig. 17.10 Edge stiffening (a) Preferred edge flange designs to increase panel stiffness; (b) Thickening
the edge flange may increase cycle times.
Fig. 17.12 Rib geometry for non-appearance surfaces.
392 Matched metal compression molding of polymer composites 17.7.7 BOSSES
17.7.9 MOLDED-IN THREADS
Projections from a plane surface of a part, called bosses, provide attachment and support for related components. They may be solid, hollow or have molded in inserts. They should also follow the guidelines for draft angles and nominal thickness (Fig. 17.13).
It is difficult to mold a thread into reinforced composites and requires highly sophisticated and costly molds and molding procedures. Molded threads should be rounded rather than sharp. Rounded threads will resist chipping and cracking and will also facilitate flow of molding material into all areas of the thread. Molded threads are usually preferred over inserts if the threaded hole diameter is over 12.7mrn (0.5 in), unless the thread is to be subjected to continual fastening and unfasteningz4. 17.7.10 MOLDED SURFACES
Nominal thickness should be maintainedthroughout part
I
Fig. 17.13 Boss design.
17.7.8 INSERTS
Inserts are objects (usually metal) which are molded into a part to facilitate repetitive fastening and unfastening of associated parts and can be provided with male or female threads. Inserts can provide bearing or bushing surfaces, electrical or other mechanical connections. Inserts should have knurls, grooves or shoulders to lock them in place and should be located parallel to the direction of mold travel.
Surfaces exhibited by the part as it comes from the mold have not been subject to any postmolding operation other than the removal of flash. The surface of the mold will reproduce itself as the surface of the part. Many different textures can be produced on the surface of a part. High gloss surfaces can be produced by highly polished molds. Draft is critical when parts are to be textured on a vertical wall. For every 0.254 mm (0.001 in) of texture depth, draft must be increased by 1". Raised or indented characters can be molded into the part. The characters should be rounded and smooth and positioned on the surfaces parallel to the parting line of the mold. 17.8 TOOLING
A good set of matched metal chromed steel tools is required for the optimum conditions when one is molding reinforced composites. Anticipated production quantities expected from the mold or the product end use, or both, should dictate the choice of steel as shown in Table 17.B5. 17.8.1 MOLD STRESSES
It is important to consider the stresses created by the flow of material at typical molding
Tooling 393 pressures from 4.13 to 8.37 MPa (600 to 1200 forced composite parts should be designed so psi). Due to unbalanced flow, narrow mold the height of any projecting mold section does sections that project from the mold surface not exceed two times the width of its base. could bend or break under such stresses. To Angular sections must not be less than 30°4 ensure sufficient strength in the mold, rein- (Figs 17.14 and 17.15).
Table 17.8 Mold steel selection5 A . Production planning volumes
Type of steel Planning volumes
Core
Cavity
5000-20 000 parts/y
AISIa-1045 steel
AISI-1045 steel
20 000-30 000 parts/y
AIS14140 forged steel prehardened to Rockwell C of 28-32
AIS14140 forged steel prehardened to Rockwell C of 28-32
Over 30 000 parts/y
AIS14140 forged steel prehardened to Rockwell C of 28-32
P-20 forged steel prehardened to Rockwell C of 28-32
100 000 parts or less
AISI-1045 steel
AISI-1045 steel
100 000-200 000 parts for mold life
AIS14140 forged steel prehardened to Rockwell C of 28-32
AIS14140 forged steel prehardened to Rockwell C of 28-32
Over 200 000 parts during mold life
AIS14140 forged steel prehardened Rockwell C of 28-32
P-20 forged steel prehardened to Rockwell C of 2&32
AISI-1045
AISI-1045 steel
AIS14140
P-20 forged steel
B. Product end use Structural items where surface appearance is not critical, such as reinforcing panels, truck front ends, etc., where molded surface quality is of secondary importance. High-quality surface appearance decorative items, such as grille opening panels, head lamp surroundings, quarter wheel opening covers, etc. where a high degree of polish is required on the outer part of cavity surface. a
American Iron and Steel Institute
394 Matched metal compression molding of polymer composites
B
Fig. 17.14 Projecting mold section; any projecting Fig. 17.15 Angular mold section: should not be mold section should not exceed two times the <30°. width of its base.
17.8.2 SHEARS All matched metal compression molds use telescoping shear edges around the perimeter of the part. The shear of the mold halves is never in contact but bypass each other as the mold closes, leaving a thin amount of flash (0.154-0.254 mm, 0.006-0.01 in). The bypassing feature allows the mold cavity to be fully filled regardless of small variations in charge weight. A minimum of 3" of draft is preferred for return flanges. A minimum of 1.0 mm (0.04 in) nominal flat clearance should be provided to keep the cavity and core halves of the mold from contacting each other. A surface normal to die draw should be provided at the edge of the part. A nominal angular tolerance of QO" and a 1.5 mm radius should be allowed for mold
building and finishing allowances Fig. 17.16). Knife edge shears are to be avoided as they create a thin mold section which can bend or break under molding pressures (Fig. 17.17). Shear edges should be flame hardened to a Rockwell C of 55-60. 17.8.3 HEELBLOCKS Compression molds should have heel blocks and wear plates suitable to withstand all lateral forces at 12.41 MPa (1800 psi) molding pressure. The heel blocks should be an integral part of the mold and flame hardened to a Rockwell C of 50-55. Bronze wear plates should be bolted opposite the heel blocks and have a minimum of 3.18 mm (1/8 in) chamfer lead in to avoid shearing off the wear plate.
Acknowledgements 395
l.0mm (0.04")
min flat
Normal to die draw +/- 200
-1I
t
+0.15mm (0.006")
flash
Fig. 17.16 Shear edge design; part design requirements to maintain sufficient tool strength at the shear edges.
17.8.5 MOLD STOPS
Fig. 17.17 Knife-edge shear should be avoided. It can be eliminated by adding a minimum flange of 5 mm (0.2 in) to the part edge.
17.8.4 GUIDE PINS
Compression molds should have leader or guide pins, the diameter of such to be a minimum of 2% of the width plus length of the mold. All guide pins should be the same height and be a minimum of 6.35 mm (0.25 in) longer than the highest point of the core. They should be chamfered for lead in and guide pin retention should be a minimum of 1.5 times the pin diameter. One guide pin should be offset to prevent misalignment of core and cavity.
Mold stops should be provided to control the mold's vertical travel. They should be flat and oil-hardened steel to a Rockwell C of 55-60. The minimum part thickness can be controlled by the mold stops. 17.8.6 SURFACE POLISH
The mold surface should be polished to the degree that is required on the surface of the part. The final 'stoning' or polishing on vertical walls should be done in draw direction. ACKNOWLEDGEMENTS
The authors would like to express their gratitude to Mr. M. Kilpinen and Mr. E. Kleese of GenCorp Automotive for their support and permission to write this manuscript. Thanks are due to Ms. Marialyce Orr for editing and proof-reading the manuscript.
396 Matched metal compression molding of polymer composites REFERENCES 1. Society of Plastics Industry Composites Institute. Semi-Annual Statistical Report, 24 August, 1993. 2. Young, R., 1969. Thermoset Matched Die Molding. In Handbook of Fiberglass and Plastics Composites, (ed. G. Lubin) pp. 391-448. New York Van Nostrand Reinhold. 3. Guillion, D. 1983. Section 1E, 38th Annual Conference of Composite Institute, The Society of the Plastics Industry, inc. 4. SMC Design Manual. 1991. Catalog Number AF-180. (a) pp. 1; (b) pp. 9; (c) pp. 12; (d) pp. 25. SMC Automotive Alliance, Society of Plastics Industry, Inc. 5. Meyer, R. 1987. (a) pp. 1-18; (b) pp. 179-204; (c) pp. 47-64. Handbook of Polyester Molding Compounds and Molding Technology. New York: Chapman & Hall. 6. Kalluar, M. 1990. Section 4F, 45th Annual Conference of Composite Institute, The Society of Plastics Industry, Inc. 7. Weaver, W.I. 1951. US Patent 2549 732. Production of Polymerized Unsaturated Resin Materials of Superior Water Resistance. 8. Frilette, V. 1951. US Patent 2 568 331. Copolymerization of Styrene and Unsaturated Resins. 9. Kia, H. 1993. (a) pp. 1-3; (b) pp. 29-48; (c) pp. 49-78; (d) pp. 95-114. In Sheet Molding Compound: Science and Technology, (ed. Kia Hamid). Cincinnati: Hanser/Gardner Publications, Inc. 10. Gupta, M.K., Hoch, D.W. and Keegan, J.F. 1987. Mod. Plast. 64:70.
11. Boening, H.V. 1964. Unsaturated Polyesters: Structure and Properties. Amsterdam: Elsevier. 12. Bruins, P.F. 1976. Unsaturated Resin Technology. New York: Gordon & Breach. 13. Anderson, T.F. and Messik, V.B. 1981. Vinyl Ester Resins. In Developments in Reinforced Plastics I. London: Applied Science. 14. Dow Chemical Company. 1975. DERAKANE Vinyl Ester Resins for Corrosion Resistance. 15. Walewski, L. and Stockton, S. 1985. Mod. Plast. 6278. 16. Edwards, H.R. 1987. Mod. Plast. 64:66. 17. Butwin, F.J. and Howes. W.C. 1986.41st Annual Conference of Composite institute, The Society of the Plastics Industry. 18. Bartkus, E.J. and Kroekel, C.H. 1970. Appl. Polymer. Symp. 15:113. 19. Atkins et al. 1976. Section 2E. 31st Annual Conference of Composite Institute, The Society of the Plastics Industry, Inc. 20. Monte, S.J. 1978. In Handbook of Fillers and Reinforcements for Plastics. New York: Van Nostrand Reinhold. 21. Mallick, P.K. 1990. In Composite Materials Technology: Processes and Properties, (ed. Mallick, P.K. and Newman, S.) pp. 25-32. New York: Oxford University Press. 22. Heimbuch, R.A. and Sanders, B.A. 1978. In Composite Materials in the Automobile Industry. American Society of Mechanical Engineers : 111-137. 23. Owens-Corning Corp. 1984. Publication 54% 12910. (a) pp. 2; (b) pp. 3. 24. Owens-Corning Corp. 1986. Publication 5-PL23920. (a) pp. 5; (b) pp. 16; (c) pp. 22.
TEXTILE PREFORMING
18
Frank K. KO and George W. Du
18.1 INTRODUCTION
resulting composite can be tailored. These fiber placement methods create textile preTextile preforming is a fiber placement forms which possess a wide spectrum of pore method utilizing textile processes prior to the geometries and pore distribution; a broad formation of composite structures. Textile prerange of structural integrity and fiber volume forms are the structural backbone of a fraction; and fiber orientation distribution as composite analogous to the structural steel well as a wide selection of formed shape and framework in a building. Starting with linear net shape capability. assemblies of fibers in continuous and/or disAs illustrated in Fig. 18.1, textile preforming crete form, these micro-fibrous structures can provides a link between raw material systems be organized into one-, two- or three-dimenand the composite product. Depending upon sional structures by means of twisting, the textile preforming method used, the range interlacing, intertwining or interlooping. By of fiber orientation and fiber volume fraction proper selection of the geometry of the fibrous of the preform will vary, subsequently affectstructures and architecture and the method of ing matrix infiltration and consolidation as placement or geometric arrangement of the well as the translation efficiency of fiber propfibers, the structural performance of the erties to composite product.
Vacuum Impregnation
K
Y I..CU
Squeeze Casting Vacuum Infiltration
Design and material system selection
Liquid Impregnation
Fig. 18.1 The role of preforms in composite processing. Handbook of Composites. Edited by S.T. Peters. Published in 1998by Chapman & Hall, London. ISBN 0 412 54020 7
Composite Product
398 Textile preforming When combined with high performance fibers, matrices and properly tailored fiber/matrix interfaces, fiber architecture promises to expand the design options for the manufacturing of tough and reliable structural composites. With an integrated network of structural cells in two- and three-dimensional arrangements, textile structures not only provide a mechanism for structural toughening of composites but also facilitate the processing of composites into net or near net shape structural parts. Considering the important role which textile preforms play in the chain of composite manufacturing processes, there is a worldwide revival of interest in the technology and science of the processing of textile preforms for composites. The serious interest in the subject of textile composites can best be illustrated in the two recent publications on textile composites, one in English (Chou and KO, 1989) and the other in Russian (Tarnpolski’i et al., 1987). These two books have been translated respectively into Russian and English. Special journal issues have been devoted to the subject of textile composites as well (for example, Jouvnal of The Textile Institute, 1990, No. 4, Textile Institute). This article intends to provide a quick reference to textile preforming from the point of view of composite processing science. To facilitate discussion, textile preforms are classified into linear, planar and three-dimensional fibrous assemblies. After a brief introduction of the processing technology of textile preforms, discussion will focus on the processing kinematics of the preforms and the geometric parameters which control two of the most important parameters in composite design: fiber volume fraction (V,) and fiber orientation (0). The role of fiber archtecture in the forming, resin infiltration and the translation of fiber properties to the composite is also discussed along with experimental evidence. For readers interested in further pursuing the technological aspect of the subject, there are several outstanding general references on industrial textiles (Kaswell, 1963;
Svedova, 1990). For more specific references one can examine the book by Goswami et al. (1977) on yarns (linear fibrous assemblies); Lord et al. (1973) on weaving; Spencer (1983) and Raz (1987) on knitting and Krcma (1971) on nonwovens. 18.1.1 THE ROLE OF TEXTILE PREFORMS
The final goal of manufacturing a composite structure is to meet design requirements including performance and cost. How successfully the goal can be met depends on the effective use of the reinforcement material and the cost of manufacturing a quality product. Preform fiber architecture plays a key role in composite manufacturing by facilitating processing steps including forming and resin infiltration. The properties of the composite will also vary depending on the fiber architecture. Fiber orientation (13)and volume fraction (V,)are key engineering parameters for textile composite formability, permeability and performance. Accordingly, after reviewing the experimental evidence of the dynamic interaction of process-structure-performance, this article will examine textile preforms through a unit cell based analysis relating the preform geometric parameters of V, and I3 and their relationship to textile preform performance.
Formability The manufacturing of composites often requires transformation of the fiber reinforcements into various structural shapes through net shape fabrication or formed shape processing. While 3-D textile preforming is more suitable for the creation of net structural shapes, 2-D textile preforms are usually formed into shapes by molding or stitching. One of the earlier studies of fabric formability in composite manufacturing (Potter, 1979) showed that total available deformation could be imposed on the fabric uniformly and that the modes of deformation are important parameters for fabric formability. Potter
In trod uction 399 demonstrated that weft knitted fabrics are significantly more conformable than biaxially woven fabrics because deformation of the knitted fabric in the axial, transverse and bias direction are 50%, 50% and 26% respectively, compared to 0%, 0% and 45% for the woven fabric. The same point can also be illustrated in a comparison of the shear resistance of biaxially and triaxially woven glass fabrics (Scardino and KO, 1981). Comparing the strain behavior of plain woven and triaxial basic fabric of similar area density (281.5 compared with 284.8 g/m2), under biaxial loading, it was found that the shear deformation of the triaxial fabrics is considerably more uniform than that of the biaxial fabrics (Fig. 18.2).As a result of this comparison, it was found that triaxial fabric is more adaptable to 3-D draw molding than biaxial fabrics made from the same yarn. In quantifying the formability of fabrics, Dow (1985) suggested that yarn slippage and low yarn jamming angles are required for fabric conformability. Accordingly, in fabric formability modeling, fiber volume fraction distribution, fiber orientation and fiber inter- (b) lacing intensity as well as the limit of geometric deformation (all of which are govm erned by the architecture created by specific f textile preforming techniques) must be considered. g
ORIENTATION 8 (degrees)
Shear strains .10
800
Y
Q)
600
W
z
- 400
Permeability Textile fluid flow permeability is an indication of how easily and uniformly a matrix can be infiltrated into the fibrous assembly. McCarthy et al. (1991)concluded that the permeability of textile preforms is mainly affected by fiber volume fraction and fiber orientation (Fig. 18.3). In the same figure, it can be noted that, for the same fiber volume fraction, ordered structures such as 0/90° woven fabrics have higher permeability than disordered structures such as discrete chopped fibers. The dependence of permeability on fiber volume fraction was also observed by Loos et
(L IL
G
2
200
I v)
'0
(4
20
40
ORENTATION
60
80
100
e (degree)
Fig. 18.2 (a) Triaxial and biaxial plain-weave fabric specimens partially deformed (50% of maximum force) in ball-burst test at same displacement of ball into fabric plane; (b) effect of test direction on the shear stiffness of biaxial fabrics at various normal stresses; (c) effect of test direction on the shear stiffness of triaxial fabrics at various normal stresses.
400
Textile preforming meability of the preforms, especially for preforms with high fiber volume fraction. The Koczeny-Carman equation was found to be adequate to provide a quantitative relationship between permeability and preform porosity.
6W4 iayers
;DO
233
Properties
Fig. 18.3 Permeability against fiber volume fraction of textile preform. 0.1
0.05
0.02
1 -
Fiber architecture plays an important role in the translation of fiber properties to the composites as well as controlling the level of matrix infiltration. The dynamic interaction of material (fiber),fiber architecture and composite properties are best illustrated by Figs. 18.5 and 18.6. In Fig. 18.5 (Dow, 1985), the in-plane tensile and shear properties of carbon/epoxy composites having a 60%)fiber volume fraction are compared for unidirectional angle ply structure, 2-D woven 3-D braided structures.
0.01 ;
0.005
-
0.002
-
0.001
-
N -
9 z Y
a
&,mu
&:
0.0005
4
n a_oooz 0.0001 ~ " ' " ' ' 0.20 0.25
'
'
I ' " " '
0.30
0.35
2
' 0.40
0.45
0.50
0.55
0.6C
Porosity, 0
Fig. 18.4 Normalized through-the-thickness permeability against porosity for multiaxial warp knit fabric preforms.
a2. (1991), as shown in Fig. 18.4. In their study of carbon multiaxial warp knit preforms, it was found that the introduction of through thickness fibers significantlyincreased the per-
Fig. 18.5 Range of values of in-plane Young's modulus, E.r and shear modulus GXYwith T-300 carbon fibers in 5200 epoxy at a volume fraction reinforcement of 0.6.
Introduction 401 e
Owoven laminates
e
a
18.1.2 CLASSIFICATION OF TEXTILE PREFORMS
AUnldirectional
There is a large family of textile preforming methods suitable for composite manufactura P ing (KO, 1989).The key criteria for the selection of textile preforms for structural composites are (a) the capability for in-plane multiaxial reinforcement, (b) through thickness reinforce00 01 02 03 04 05 ment and (c) the capability for formed shape Void fraction and/or net shape manufacturing. Depending Fig. 18.6 Effect of fiber architecture on the flexural on the processing and end use requirements, some or all of these features are required. strength of SiC/SiC composites. On the basis of structural integrity and fiber It can be seen that with each reinforcement linearity and continuity, fiber architecture can group, fiber orientation has a significant influ- be classified into four categories: discrete; conence on the overall performance of the tinuous; planar interlaced (2-D); and fully composite. Within the envelope of in-plane integrated (3-D) structures. In Table 18.1 the properties of unidirectional reinforcement lies nature of the various levels of fiber architeca broad range of properties for 2-D and 3-D ture is summarized (Scardino, 1989). A discrete fiber system such as a whisker or textile composites. For very brittle composites such as the fiber mat has no material continuity; the orienSiC/SiC composites produced by the chemical tation of the fibers is difficult to control vapor infiltration (CVI) process, the effect of precisely, although some aligned discrete fiber fiber architecture on the level of infiltration is systems have recently been introduced. The quite evident. Comparing the flexural strength structural integrity of the fibrous preform is of short fiber mat, unidirectional tape, woven derived mainly from inter-fiber friction. The laminates and 3-D braided reinforced ceramic strength translation efficiency, or the fraction matrix composites, one can see that a 3-D fiber of fiber strength translated to the non-aligned architecture produces structures superior to fibrous assembly of the reinforcement system, others because of the three-dimensional inter- is quite low. The second category of fiber architecture is connected network of Sic fibers. The unidirectional composite has the highest vari- the continuous filament, or unidirectional ations of void factor and unidirectional system. This architecture has the highest level strength. It is clear in Fig. 18.6 that the perfor- of fiber continuity and linearity and consemance of composites depends not only on the quently has the highest level of property volume fraction but also the fiber orientation translation efficiency and is very suitable for filament wound and angle ply tape lay-up of the preform (KOet aZ., 1988a). P
(OO)
Table 18.1 Fiber architecture for composites
Level
Reinforcement system
Textile construction
I I1 I11 IV
Discrete Linear Laminar Integrated
Chopped fiber Filament yarn Simple fabric Advanced fabric
~~~~-~ ~~
~
_
Fiber length
Fiber orientation
Fiber entanglement
Discontinuous Continuous Continuous Continuous
Uncontrolled Linear Planar 3-D
None None Planar 3-D
_
402 Textile preforming structures. The drawback of this fiber architecture is its intra- and interlaminar weakness due to the lack of in-plane and out-of-plane yarn interlacing. A third category of fiber reinforcement is the planar interlaced and interlooped system. Although the intra-laminar failure problem associated with the continuous filament system is addressed with this fiber architecture, the interlaminar strength is limited by the matrix strength due to the lack of through thickness fiber reinforcement. The fully integrated system forms the fourth category of fiber architecture wherein the fibers are oriented in various in-plane and out-of-plane directions. With the continuous filament yarn, a three dimensional network of yarn bundles is formed in an integral manner. The most attractive feature of the integrated structure is the additional reinforcement in the through-thickness direction which makes the composite virtually delamination-free. Another interesting aspect of many of the fully integrated structures such as 3-D woven, knits and braids is their ability to assume complex structural shapes. Another way of classifying textile preforms is based on the fabric formation techniques: through fiber entanglement or yam twisting, interlacing, interlooping, intertwining or multiaxial placement. While most textile preforms are converted from fiber to yarn or yarn to fabric structures, some preforms, such as fiber felts, are converted directly from fiber to fabric. In Table 18.2, the four basic yarn-to-fabric formation techniques are compared.
While weaving, braiding and knitting can produce planar or 3-D structures, nonwoven fabrics can be a 2-D planar system with random or organized fiber orientation, as well as the orthogonal 3-D system. The 2-D and 3-D fabrics are distinguished by yarn orientation distribution and the number of yarn diameters in the thickness direction. A 2-D fabric consists of two to three yarn diameters in the thickness direction with fibers oriented in the x-y plane. A 3-D fabric, consisting of three or more yams in the thickness direction, is a fibrous network wherein yarns pass from surface to surface of the fabric in all three directions.
18.2 YARN PREPARATION FOR TEXTILE PREFORMING
18.2.1 CLASSIFICATION OF YARNS
In addition to preforming methods and parameters, the physical properties of textile preforms, such as thickness, fabric tightness, fiber orientation, etc., are also affected by the characteristics of yarns. Linear fibrous assemblies can be composed of filaments of discrete (staple yarns) or continuous lengths (filament yarns). Staple yarns are held together by an appropriate level of twist, whereas filament yarns may or may not have twist. Larger filament bundles that have little or no twist are called rovings; yams are usually smaller than rovings and some level of twist is generally added. The majority of high performance yarns are continuous filament yams having single or multiple strands.
Table 18.2 A comparison of fabric formation techniques ~
Preforming technology
Yarn introduction direction
~~~
Formation technique
Weaving
Two ( O o / 9 O 0 ) (warp and fill)
Interlacing (by selective insertion of 90" yarns into 0" yarn system)
Nonwoven
Three or more (orthogonal)
Mutual fiber placement
Knitting
One (0" or 90") (warp or fill)
Interlooping (by drawing loops of yams over previous loops)
Braiding
One (machine direction)
Intertwining (position displacement)
Yarn preparation for textile preforming 403 18.2.2 YARN GEOMETRY
The geometric parameters which describe a linear fiber assembly include: shape of bundle cross-section, number of fibers in the bundle, dimension and shape of fiber cross-section, bundle twist level, degree of fiber migration in the radial direction and fraction of interfiber packing. Usually, the fiber bundles are assumed to be circular in cross-section, but in reality, the fibers or filaments can be packed in to various shapes. Most of the engineering fibers, such as glass and carbon, have a circular or near-circular cross-section with a constant diameter. For some ceramic fibers, such as Nextel, the fibers have an ellipse shape and varying dimensions. The filament bundles used for composites have a small amount of twist, usually less than 4 tpi, the bundle surface twisting angle is small. The geometry of interfiber packing in fiber bundles has been studied by a number of researchers primarily for textile applications (Hearle et al., 1969). Three basic idealized forms of circular fiber packing were identified: open-packing, in which the fibers are arranged in concentric layers (Fig. 18.7(a));square-packing, in which the fibers are enclosed by a square (Fig. 18.7(b)); and close-packing, in which the fibers are arranged in a hexagonal pattern (Fig. 18.7(c)). In open-packed bundles the fiber volume fraction, defined as fiber to bundle area ratio, has been computed as a function of the number of fibers. If the outer ring is completely filled and the fibers are circular, the fiber
volume fraction can be shown to equal (Du, Popper and Chou, 1991) Yopen
= 1-
3N1(N1- 1) + 1 (2N1- 1)’
(18.1)
where N , is the number of rings and its relationship to the number of fibers, N,, is given by 1 N, = 2 +
1 {[-41 + -(2Nf 3
- I)]
(18.2)
For large numbers of fibers the fiber volume fraction approaches 0.75. In squared packed bundles, the fibers are arranged in a square array. For any number of circular fibers, if the outer layer is completely filled, the fiber volume fraction can be shown to equal the area ratio of a circle to an enclosing square:
x Vfsquarr = 4
= 0.785
(18.3)
Similar to square packing, the fiber volume fraction of close packed bundle is equal to the area ratio of a circle to an enclosing hexagon: 0.907
(18.4)
The level of bundle fiber volume fraction predicted by the above models assumed circular, square and hexagonal fiber bundles for open, square and close packing, respectively. However, they apply equally well to other shapes if the number of fibers is sufficiently large.
Fig. 18.7 Idealized fiber packings: (a) open packing; (b) square packing; (c) close packing.
404
Textile preforming (surface helix angle) and the twist level, as shown in Fig. 18.3. Clearly, for a given twist inserted to the fiber bundle, as fiber orientation angle increases, yarn diameter increases (as indicated by equation (18.6)) whereas the fiber volume fraction decreases as can be seen in Fig. 18.9, which is useful in determining the twist level of fiber bundles. For example, to obtain a fiber volume fraction (or fiber packing fraction) 0.8 and a fiber orientation angle lo", twist level of 3 tpi should be used for the 12K, 7 pm fiber diameter carbon yarns.
Fig. 18.8 Geometry of twisted yarn.
For twisted fiber bundles, the fibers are no longer aligned along the bundle axis. Instead, the fibers assume a helix configuration within the bundle, as shown in Fig. 18.8. The fibers in different radial layers of the bundle have a different helix angle:
Bi = tan+ [n(D,- d)T]
(18.5)
where d is the fiber diameter, T is twist level and Di is the diameter of the fiber layer. Apparently, fibers at the outer layer have a maximum orientation (helix) angle: 8 = tan-' [n(D- d)T]
0 10 20 30 40 50 60 70 80 90 9 (")
Fig. 18.9 Relationship of fiber volume fraction to (18.6) fiber orientation at various twist levels.
where D is the diameter of the bundle. According to Hearle (1969), yarn diameter is related to the number of filaments ( n ) in the yarn and the packing fraction of the fibers ( K ) in the following relationship:
18.3 WEAVING
18.3.1 PROCESSING TECHNOLOGY
(18.7) Weaving, which is the interlacing of two sets of yams usually at a right angle to each other, Viewing a yarn bundle as an individual pre- requires holding one set of yams in parallel form, its fiber volume fraction is actually equal rows and passing another set over and under to its fiber packing fraction, i.e. V , = IC. the first set. The set of lengthwise yarns is Combining equations (18.6) and (18.7), we called warp and the set of crosswise yarns is the fill. The simplest two-hamess loom is have sketched in Fig. 18.10 and may be used for -2 (18.8) making biaxial plain weave fabrics shown in Fig. 18.11(a).For more complex weaves such For a 12K carbon yarn with 7 p m fiber diame- as twill and satin as illustrated in Figs 18.11@). ter, one can establish the relationship between and (c), looms with more than two harnesses its fiber volume fraction, fiber orientation are required.
Weaving 405 harnesses are lifted. Fabrics with as many as 17 layers have been successfully woven with this method.
beam
Shuttle
Reed
Fig. 18-10A simple two-harness weaving loom (Smith and Block, 1982).
In a novel departure from standard weaving methods, triaxial weaving interlaces two warp yarns and one filling yarn at 60" angles as shown in Fig. 18.12 (Dow et al., 1970).The triaxial fabric has excellent dimensional stability and is currently used almost exclusively for industrial products. Three-dimensional woven fabrics are produced principally by the multiple warp weaving method, long used for the manufacturing of double cloth and triple cloths for bags, webbings and carpets. A typical setup of a multiwarp weaving loom is shown in Fig. 18.13.The number of layers of yarns in the fabrics is governed by a special shedding mechanism which controls the height that the
Fig. 18.12 Structural geometry of triaxially woven fabrics (basic weave).
p\
;B \
a:
\
Fig. 18.13 Setup of a multiwarp weaving loom.
Fig. 18.11 Structural geometry of biaxially woven fabrics: (a) plain weave; (b)twill weave; (c) satin weave.
406 Textile preforming 18.3.2 STRUCTURAL GEOMETRY
Biaxial weaves consist of 0" and 90" yarns interlaced in various repeating patterns or topological unit cells. The three basic weave geometries from which many other patterns evolve are the plane, satin and twill weave. A schematic diagram for the various views of these three basic weaves are shown in Fig. 18.11.These three fabrics are distinguished by their frequency of yarn interlacing and the linearity of the yarn segments. The plane weave has the highest frequency of yarn interlacing whereas the satin weave has the least number of yarn interlacing, with the twill weave somewhere in between. Accordingly, the plane weave has a higher level of structural integrity and greater ductility due to the crimp geometry produced by yarn interlacing. On the other hand, the satin weave has the highest level of fiber to fabric strength and modulus translation efficiency due to the low level of yarn interlacing and yarn linearity. The low level of yarn integration in satin weave also allows freedom of yarn mobility which contributes to higher fiber packing density and consequently higher level of fiber volume fraction.
Fig. 18.14 Structural geometry of various 3-D woven fabrics: (a) solid orthogonal panel; (b) variable thickness solid panel; (c) rectangular core structure; (d) triangular core structure; (e) angle interlock.
Triaxial weave has 90+60" yarns oriented in one plane, resulting in a high level of in-plane shear resistance (Dow et al., 1970).High levels of isotropy and dimensional stability can be achieved with triaxial weave at low fiber volume fraction. Figure 18.12 shows a schematic diagram of triaxial weave geometry. Using the multiwarp weaving method, various fiber architectures can be produced including solid orthogonal panels, variable thickness solid panels and core structures simulating a box beam, or truss-like structure as illustrated in Fig. 18.14. Furthermore, by proper manipulation of the warp yarns, as exemplified by the angle interlock structure, the through-thickness yarns can be organized into a diagonal pattern, as shown in Fig. 18.14(e). One limitation of the multiwarp weaving method is the difficulty of introducing yarns in the bias direction as in the triaxial weaving or circular weaving process. However, this is now being addressed by a modification of triaxial weaving techniques (Dow, 1989). As illustrated in Fig. 18.15 (Pastore and Cai, 1990a), there are four basic components to a generalized three-dimensional woven fabric
Weaving 407 geometry: warp, web, fill and surface weave yarns. Warp yarns are the system of yarns which run in the machine direction and have no crimp. These are also called ’stuffer’yarns or ’longitudinals’. Because of their very low crimp, these yarns provide the primary strength and stiffness in the longitudinal (x) direction of the material. Web yarns run in the machine direction and provide the interlacing necessary for fabric integrity. These yarns contain crimp in the through thickness direction, providing the z-directional properties of the system. These yarns are sometimes called ’weavers’. The ‘weave angle’ of the web yarns (6) refers to the angle of orientation of the web yarn with respect to the warp direction. Fill yarns are perpendicular to machine direction and interlace with the web yarns. These yarns are sometimes called ’picks’. These yarns also possess crimp in the through thickness direction, but this crimp is negligible compared to that of the warp yams for these fabric systems. These yarns provide the transverse (y) directional properties of the composite system. Surface weave yarns run in the machine direction and form what is essentially a two-dimensional weave on the surface of the fabric. Surface weave yarns are incorporated into the structure when the web yarns are insufficient to provide a smooth surface on the face and back of the cloth. These yarns experience crimp in the through thickness direction.
When surface weave yarns are employed in the fabric, there are two yarns for every warp plane of the fabric. This system of yarns contributes the least to the mechanical properties of the composite. 18.3.3 DESIGN METHODOLOGY
Figure 18.16 gives the unit cell geometry for plain biaxial weave, as proposed by Dow and Ramnath (1987). In their analysis, Dow and Ramnath assumed circular yarn cross-section, the same yam diameter and pitch length for both fill and warp yarns. The expression of the fiber volume fraction was derived:
I
2-+46
(18.9)
where K is the fiber packing fraction, d is the yarn diameter, L is the pitch length, T is the fabric thickness, I is the dimension shown in Fig. 18.16. The yarn inclination angle to the fabric plane, 6, is given by:
Surface weaver
Weaver or web yarn
Pick or filling yarn Warp yarn
Fig. 18.15 Schematic illustration of generalized 3-D woven fabric projected to the x-z (fabric length-thickness) plane.
408 Textile preforming lengths for fill and warp yarns can be analyzed. Based on the structural geometry shown in Fig. 18.15, the orientations and volumetric distributions of all yarns in the 3-D weave can be calculated also using the unit cell method. Detailed analysis is given by KO and Du (1992).
__--V
t .+I
I
10
09 08
> 06 SCCTION
05
-
04
.-.
03
Fig. 18.16 Unit cell geometry of plain weave.
02 01
00
The fabric thickness is very close to two yarn diameter, i.e.
0
10 20
30 40 50 60 70 80 90 0()
T=2d
(18'11)
Fig. 18.17 Relationship of fiber volume fraction to fiber orientation for plain weave.
1 - 1 d tan8
(18.12)
18.4 KNITTING
and approximately:
Equation (18.9) is then simplified to
18.4.1 PROCESSING TECHNOLOGY
Knitting is the interlocking of one or more yarns through a series of loops (also called stitches). The lengthwise columns of stitches corresponding to the warp in woven fabrics are called wales; the crosswise rows of stitches Figure 18.17 plots the fiber volume fraction corresponding to the filling are known as against the yarn inclination angle. It can be courses. Knitted structures can be classified by seen that as the inclination angle increases, the basic loop formation mechanism into weft pitch length becomes longer which results in a knits and warp knits. In weft knitting, as lower fiber volume fraction. The woven fabric shown in Fig. 18.18(a),yam feeding and loop has the tightest structure at the inclination formation occur at each needle in succession angle of 60" (when Lld = 3 in equation (18.10)). along the wale direction and all the courses of In this calculation, the fiber packing fraction K loops are composed of single strands of yam. is assumed to be 0.8. In warp knitting, there is a simultaneous yarnThe above analysis is given only for the sim- feeding and loop-forming action occurring at plest of woven structures. Different weave every needle and all the wales of loops are patterns, non-circular yarn cross-sectional composed of single strands of yarn as illusshape, different yarn dimensions and pitch trated in Fig. 18.18(b).
Knitting 409
J
’Djrection
of knitting
Fig. 18.18 Yarn feeding and loop formation: (a) weft knitting; (b) warp knitting (Spencer, 1983).
technology can be found in Spencer (1983)and Raz (1987). Knitted 3-D fabrics are produced either by weft or warp knitting. An example of a weft knit is the near net shape structure knitted under computer control by the Pressure Foot@ process (Williams, 1978). In a collapsed form this preform has been used for carbon-carbon aircraft brakes. While weft knitted structures have applications in limited areas, multiaxial warp knit (MWK) 3-D structures are more promising and have undergone a great deal more development in recent years. Schematic of a MWK LIBA system is given in Fig. 18.20, in which up to six layers of insertion yarns plus one layer of non-woven can be stitched together.
Stitch (loop) formation is similar in both weft and warp knitting. The formation of the stitches in a single wale is illustrated in Fig. 18.19. In Step 1, the needle rises through loop A from its lowest position; in Step 2, yarn slips under the tip of the needle and onto the stem; in Step 3, ascending hook catches the new yarn at the top of its rise and begins to descend; in Step 4, the new yarn slips under the tip and into the hook; in Step 5, the needle moves down until the tip slides under loop A and the hook pulls the new loop through. After the completion of five steps, loop B is formed and the process is repeated. In a knitting operation, each of the needles is controlled by a cam to rise and fall in synchronization with the other needles. Detailed description of the knitting ,Hook
Ti
1
2
3
Fig. 18.19 Stitch formation in knitting machines (Smith and Block, 1982).
410 Textile preforming
Fig. 18.20 Multiaxial warp knit with four layers ( O O , 90" and &) of inserted yarns and (a) chain stitch or (b) tricot stitch.
18.4.2 STRUCTURAL GEOMETRY
ture for the incorporation of 0" and/or 90" insertion yarns. Knitted fabrics are traditionally identified The MWK fabric system consists of warp (O"), with socks, underwear and sweaters. In the yarns held together by search for methods to reduce composite man- weft (90") and bias (d) a chain or tricot stitch through the thickness of ufacturing costs, textile preforms including the fabric, as illustrated in Fig. 18.21. knitted structures are receiving increased Theoretically, the MWK can be made to as interest in the composite industry. While conmany layers of multiaxial yams as needed, but formability and productivity are obvious current commercially available machines only attributes for knitted preforms, the availability of a broad range of micro- and macrostructural allow four layers (the Mayer system) of 0", 90", geometries has only recently been recognized. +O and 4 insertion yams, or six layers (the The non-linearity of knitting loops, severe LIBA system) of 2(90"), 0", 2(+8) and 2 ( 4 ) bending of yams during the knitting process insertion yarns to be stitched together. All layand limited fiber packing density resulting in ers of insertion yarns are placed in perfect the formation of resin pockets within a knit- order each on top of the other in the knitting ting loop prevent kmts from being considered process. Each layer shows the uniformity of the uncrimped parallel yams. The insertion yarns for structural applications. The development of technology for the usually possess a much higher linear density directional insertion of linear yarns in weft than the stitch yarns and are therefore the and warp knits greatly enhances opportunities major load bearing component of the fabric. for knitted preforms for conformable structural composites by combining the 18.4.3 DESIGN METHODOLOGY conformable foundation knit structure with directional reinforcement. As shown in Fig. Similar to the 2-D woven fabrics, the unit cells 18.20, sewing threads (high twist yams) or for the knitted structures are also different, very fine yams are used to form a base struc- depending on the knit constructions such as
Knitting 411
Fig. 18.21 Multiaxial warp knit LIBA system.
stitch patterns and laid-in insertions. To illustrate the use of the unit cell method for relating fiber volume fraction, yarn orientation and processing variables, a plain weft knit as shown in Fig. 18.18(a) is selected as an example. The unit cell geometry identified for the plain weft knit is shown in Fig. 18.22, having a dimension of x (course width), y (half wale width) and z (fabric thickness).
illustrated in Fig. 18.23(b).For untwisted fiber bundles under compression applied during preforming or composite processing, they have a ribbon-like cross-section similar to a race-track with a width-to-thickness aspect ratio off > 1 as illustrated in Fig. 18.23(a).For composite applications, untwisted fiber bundles are usually used in knitting, which have an aspect ratio f slightly larger than 1 at the off-machine state. To increase the fiber volume fraction for knitted structures, very high pressure will be required to reduce the knit thickness. Under the compression status, the yarn aspect ratio f can increase to as high as 12 for untwisted bundles, provided that there are no restrictions applied to yarn edges.
Fig. 18.22 Unit cell geometry of plain knitted struc-
ture. In traditional textile fabric manufacturing, highly twisted fiber bundles are used. These materials can maintain a circular shape with a width-to-thickness aspect ratio of f = 1, as
(a)
f =w/t
1
(b)
f =1
Fig. 18.23 Idealized yarn cross-sections: (a) racetrack cross-section with width-to-thickness aspect ratio f > 1; (b) race-track shape becomes circular whenf = 1.
412 Textile preforming
In this analysis, the knit thickness is assumed to be approximately equal to two yarn thickness (t)for computational purposes, i.e.
z = 2t
(18.14)
The yarn orientation angle (e), which is the angle made by the fabric axis (in x direction) and the yarn path projected to the fabric surface plane ( x - y), is given by: (18.15)
The fiber volume fraction (V,),which is defined as the ratio of volume of total fibers to the overall composite volume, can be derived as:
Y (18.16) 1+tan where k is the fiber packing fraction within yarn bundles and a is the shape correction factor defined as:
relative course width (x/w), relative half wale and yarn aspect ratio (f)under width (y/w) compression is depicted in Fig. 18.24, using the geometric model developed. In the calculation, we use the fiber packing fraction k = 0.8, which is within the range for tightly packed yarn bundles according to experimental observation. Also, to show the processing window of fiber volume fraction in highest region, one can assume yam jamming in the course (x) direction, i.e. x/w = 3 according to equation (18.18). As can be seen from Fig. 18.24, the fiber volume fraction V , decreases with the increase in relative half wale width in the range of y/w = 2-10. When y/w is beyond 10, the fiber volume fraction slightly increases and soon approaches a constant with the increase in relative half wale width. The wale width cannot be smaller than 4 yarn widths, or y/w 2 2 as given by equation (18.19). Knitted yarns have an aspect ratio f = 1 at free-stress status (as made off-machine) and the fiber volume fraction for the knitted preform has a minimum value. Figure 18.24 shows that, for the plain weft knit at its tightest possible structure (x/w = 3, y/w = 2 and k = OB), its maximum fiber volume fraction is only about 0.274. To increase the fiber volume fraction, a compression in the fabric thickness direction is necessary. The effect of the compression is the increase in yam aspect ratio (i.e. yams within the knitted structure become wider in x-y plane
(18.17)
060
The limiting geometry of the knitted structure due to yarn jamming is governed by: X W
-23
(18.18)
020
P LL
f = 1 (circular yarn without compressior 0.10
1
W
2
2
(18.19)
The processing window of fiber volume fraction for knitted structures within the possible ranges of key processing parameters, such as
ooo 10
1000
100
Ylw
Fig. 18-24 Processing window of fiber volume frattion for the plain knitted structure.
Braiding 413 but thinner in z direction).As a result, the yarn coverage over the fabric increases, whereas the volume of the preform decreases due to the decrease in fabric thickness. These two factors, the increased yarn coverage and decreased preform volume, raise the fiber volume fraction to a much higher level. As shown in Fig. 18.24, at a maximum aspect ratio f = 12, the fiber volume fraction can be as high as 0.475. A series of studies on the technology, structure and properties of the MWK preforms and composites have been reported by KO and his co-workers (1980,1982,1985,1986,198813).In a recent study, a unit cell based geometric model of the four-layer MWK structure as shown in Fig. 18.21was developed by Du and KO (1992). Based on the experimental observations, the unit cell geometry of the MWK fabric is identified and a geometric model is developed relating the fiber volume fraction and fiber orientation in terms of structural and processing parameters.
18.5 BRAIDING
18.5.1 PROCESSING TECHNOLOGY
Braiding is an old textile technology, traditionally used for the manufacture of a wide variety of linear products ranging from cables, electrical insulators and shoelaces to surgical sutures. Recognizing the high level of conformability and the damage resistance capability of braided structures, the composites industry had found structural applications for braided composites ranging from rocket launchers to automotive parts to aircraft structures. Two-dimensional braided structures are intertwined fibrous structures capable of forming structures with 0" and & fiber orientation. Although 2-D braids can be fabricated in tape form, the majority of braided structures are fabricated with a tubular geometry. Thickness is built up by overbraiding previously braided layers similar to a ply lay-up process. Braiding can take place vertically orhorizontally, but a majority of the composite braiders are horizon-
tal. A schematic of a horizontal braider is shown in Fig. 18.25.Although braiding is similar to filament winding in many ways, the major difference between braiding and filament winding is that braids are interlaced structures having as many as 144 or more interlacing per braiding cycle (or pick). Three-dimensional braiding technology is an extension of 2-D braiding technology in which the fabric is constructed by the intertwining or orthogonal interlacing of yarns to form an integral structure through position displacement. A unique feature of 3-D braids is their ability to provide through the thickness reinforcement of composites as well as their ready adaptability to the fabrication of a wide range of complex shapes ranging from solid rods to I-beams to thick-walled rocket nozzles. Three-dimensional braids have been produced on traditional Maypole machines for ropes and packings in solid, circular or square cross-sections. The yarn carrier movement is activated in a restricted fashion by horn gears. A 3-D cylindrical braiding machine of this form was recently introduced by Albany with some modification that the yarn carriers do not move through all the layers (Brookstein, 1991). 3-D braiding processes without using the horn gears, including Track and Column (Brown et al., 1988) and 2-Step (Popper and McConnell, 1987),have been developed since the late 1960s in the search for multidirectionally reinforced composites for aerospace applications. A generalized schematic of a 3-D braiding process is shown in Fig. 18.26. Axial yarns, if present in a particular braid, are fed directly Axial yarns,
/Carrier track
Fig. 18.25 Schematic of tubular braider with gantry system.
414 Textile pyeforrning
;;urbanism
,
Forming point
-
Convergence point
Fig. 18.26 Schematic of a generalized 3-D braider.
into the structure from packages located below the track plate. Braiding yarns are fed from bobbins mounted on carriers that move on the track plate. The pattern produced by the motion of the braiders relative to each other and the axial yarns establish the type of braid being formed, as well as the microstructure.
Track and column braiding is the most popular process in the manufacturing of 3-D braided preforms. The mechanism of these braiding methods differs from the traditional horn gear method only in the way the carriers are displaced to create the final braid geometry.Figure 18.27(a)shows a basic loom setup in a rectangular configuration. The carriers are arranged in tracks and columns to form the required shape and additional carriers are added to the outside of the array in alternating locations. Four steps of motion are imposed to the tracks and columns during a complete braiding cycle, resulting in the alternate x and y displacement of yam carriers, as shown in Fig. 18.27(b-e).The formation of shapes, such as T-beam and I-beam, is accomplished by proper positioning of the carriers and the joining of various rectangular groups through selected carrier movements. The track and column braiding machine can also be used to create 2-step braids and other similar 3-D structures by simply adding a certain number of axial yarns and removing most of the braiding yarns (Du and KO, 1993a).
Y
Track direction
Fig. 18.27 Formation of a rectangular 3-D track and column braid, using 4 tracks, 8 columns and 1 x 1 braiding pattern. (a) Initial loom setup; (b) Step 1: tracks move horizontally; odd tracks move to left and even tracks move to right; (c) Step 2: columns move vertically; odd columns move down and even columns move up; (d) Step 3: tracks move horizontally; odd tracks move to right and even tracks move to left; (e) Step 4: columns move vertically; odd columns move up and even columns move down.
Braiding 415 18.5.2 STRUCTURAL GEOMETRY
As with woven fabric, braids can be formed with different yarn interlacing patterns by simply changing relative position of carriers on the track ring. If one bias yarn continuously passes over one yarn and then under one yarn of the opposing group, the pattern is designated as 1/1 braid, or diamond braid as generally recognized. Other simple interlacing patterns in common use include 2/2,3/3,2/1 and 3/1 braids. Figure 18.28 shows the pattern of 2/2 braid with axial insertion. Among all these patterns, the 2/2 braid is the most popular and has been referred to as regular, standard, plain or flat braid. The path of axial yarns is independent of braid interlacing patterns, they are always over one group of bias yarns, but under the opposite group. The formation of shape and fiber architecture are illustrated in Fig. 18.29 which depicts the process of braiding over an axisymmetric shape of revolution. Braiding angle can range
from 5" in almost parallel yarn braid to approximately 85" in a hoop yarn braid, depending on the mandrel dimension, the machine speed ratio and the convergence length (Du et al., 1990). The 2-D braid can be defined as a fabric which consists of only two layers of bias yarns interlaced with each other. In 3-D braided structures, at least three layers of bias yarns go through the thickness in a zig-zag manner along the diagonal direction. Similar to the 2-D structure, longitudinal yarns can be incorporated in the 3-D braid for the enhancement of stiffness and strength in the length direction. Regardless of the difference in the carrier propelling mechanism, there are basically two types of 3-D braiding looms: rectangular and circular. The former is usually used to fabricate solid structures such as panel, I- and T-beam etc. and the latter for making thick wall tubular structures. Figure 18.30(a) shows a schematic of a 3-D braided slab.
Fig. 18.28 Yarn structure in 2-D braid: braiding yarns at & to braid axis, optional axial yarns at 0" to braid axis.
Fig. 18.29 Braid formation over a shaped mandrel.
Fig. 18.30 3-D braided solid slab (and its cross-section as seen on SEM).
416
Textile preforming
18.5.3 DESIGN METHODOLOGY
The unit cell geometry of 2-step braids has been reported by Du et al. (1991). Based on experimental observations, diamond and ribbon shapes for the axial and braider yarns, respectively, are assumed in their analysis. The unit cell was defined; pitch length and percentage of braider yarns were identified as key process parameters which control the braid microstructure and the jamming criterion for the 2-step braid was given. The traditional approach used in modeling 3-D braided composites is to artificially define a unit cell geometry for a 3-D braided structure without providing any relationship between processing variables and geometric parameters. All fibers in the unit cell are assumed to incline in four different diagonal directions, as well as along the longitudinal direction, if any. Fiber volume fraction is assumed to be either known or measured. The approach used in geometric modeling of textile structures is to first determine the dimension, shape and fiber architecture of the unit cell based on process and structural analysis; using the unit cell geometry identified, the relationship between processing variables and key geometric parameters can readily be established. The key geometric parameters of 3-D braids (which affect reinforcement capability and composite processability) include braider orientation, total fiber volume fraction, volume fraction of inter-yarn void and axial fiber percentage of total fibers. Although there are only two simple process parameters adjustable to control the microstructure of 3-D braids (speed ratio between braiding and take-up and linear density ratio of braider and axial yarns), the process-structure model of 3-D braid is complicated. Normally, yarn bundles consisting of numerous continuous filaments are used for fabric preforms, thus, the fabric microstructure has three levels: geometry of interfiber packing in the yarn bundle (fiber level),
cross-section of yarn bundles in the fabric (yam level) and orientation and distribution of fibers in the 3-D network (fabric level). The unit-cell technique is commonly used to establish the geometric relation. In most of 2-D fabrics a unit cell geometry is readily identified, but in complex 3-D fabrics it can be very difficult to define. The fiber volume fraction of a 3-D fabric depends on the level to which yarns pack against each other in the structure and the level to which fibers pack against each other in a yarn, as illustrated in Fig. 18.7. In addition to the level of packing fraction, the fibers also establish the yarn cross-sectional shape, i.e. yarn packing in fabrics. This shape plays a very significant role in determining how many fibers can be packed into a fabric. One good example is the yarn packing in 2-step braided preforms (Du et al., 1991). Due to the use of untwisted fiber bundles and high braiding tensions, cross-section of axial yarns in the 2step braid is deformed to prismatic shapes, giving most the compact yarn packing within the braided structure. For the track-and-column braids, the braiding tensions are lower compared to the 2-step braids and the crosssections of yarns actually have a polygonal shape. The microgeometric model for the trackand-column braid has been investigated by many researchers since the early 1980s (Pastore et al., 1990b; Li et al., 1990). The most recent one was given by Du and KO (1993a), which does not only relate geometric parameters and processing variables but also provides limiting braid geometry due to yarn jamming. In their analysis, the yarns are assumed as rigid circular rods. This assumption is valid when braiding at high yarn tensions. When low yarn tensions are used, yarn crimp will be introduced during braiding or during postpreforming processing due to distortion. This yarn waviness (crimp) may increase the fiber volume fraction of the braid with the sacrifice of directional reinforcing efficiency.
Braiding 417 Figure 18.31shows an idealized braid cross- where IC is the fiber packing fraction (fiber-tosection cut longitudinally at a 45" angle to the yarn area ratio). Due to the bulky fiber and braid surface. There are four groups of yarns nonlinear crimp nature, it is difficult to fabriinclined at angle 8 with the braid axis (z direc- cate the braid with tightest structure. In tion) in different directions; the yarns in each practice, the yarn orientation angle (braid group are parallel to each other within a spe- angle) is determined from the yarn diameter cific plane. Two groups of yarns are parallel to (d) and braid pitch length (kZ).The fiber volthe XI-z plane; the other two are parallel to the ume fraction is controlled by the braiding y'-z plane. The cutting plane is so selected that angle and the braid tightness factor. The govit cuts through the diameter of a group of yarns. erning equations are given below:
8
e
= sin-1
{((k,/d)2 +
4)
(kZ2 2d) (18.21)
where is the fabric tightness factor, which is within the range of 0 to ~ / 4This . tightness factor must be so selected that the required fiber volume fraction is achieved and also that the over-jamming condition is avoided. Figure 18.32 shows the V , 4 relationship prior to and at the jamming condition, based on the governing equations. The fiber packing fraction, K, is assumed as 0.785. As can be seen, there are three regions of fiber volume fraction. The upper region cannot be achieved due to the impossible fiber packing in a yarn bundle. Jamming occurs when the highest
A
C Fig. 18.31 Braid cross-section cut longitudinally at a 45" angle to the braid surface by the Y-y plane ABCD. z is the braid length direction.
The braid has the tightest structure when each yarn is in contact with all its neighboring yarns, in other words, the yarns are jammed against each other. At the jamming condition, the fiber volume fraction V ,as a function of the braid angle 8 was derived by Du and KO (1993a):
v,=-K2K
cos8 1 + cos%
(18.20)
1
0,51= 0.4 0.3 0.2 0.1
0.04 0
I
10
.
,
20
.
Y
,
30
40
50
60
70
80
90
0 (")
Fig. 18.32 Relationship of fiber volume fraction to braiding angle for various tightness factors. Fiber packing fraction K is assumed to be 0.785.
418 Textile preforming braiding angle is reached for a given fabric tightness factor q. The non-shaded region is the working window for a variety of V f 4 combinations. Clearly, for a given fabric tightness, the higher braiding angle gives higher fiber volume fraction and for a fixed braiding angle, the fiber volume fraction is greater at higher tightness factors. Theoretically, the 2-D braid can be considered as a single layer of 3-D braid. For prepreg or tape braiding without much change in yarn width and for the braiding of structures with constant cross-sections, the V f 4 relation is simple. For braiding of dry tows and structures with variable cross-sections wherein a dynamic interaction of the braiding machine and tow geometry takes place, there is a need for a more general representation of the kinematics of the braiding processes which allows for the tow width to vary over a limiting geometry. Two mathematical models have been developed, the first is the kinematics model (Du et al., 1990) which provides the relationship between the braiding angle and the braiding process parameters and the other is the unit cell model (Du and KO, 199313)which relates braiding angle to yarn geometry to predict fiber volume fractions V, along both the braiding and axial directions.
Fig. 18.33 Schematic of the Novoltexa process.
thickness fiber reinforcement. For illustration purposes, our analysis of fiber volume fraction distribution will be focused on orthogonal nonwoven 3-D fabrics. While woven 3-D fabrics have a long history of development and is clearly a product of the textile industry, the class of orthogonal nonwoven 3-D fabrics is a product of the twentieth century, developed in the aerospace industry for specific composite applications. Pioneered by aerospace companies such as 18.6 NONWOVEN General Electric and AVCO, the nonwoven 318.6.1 PROCESSING TECHNOLOGY D fabric technology was developed further by Nonwoven structures are fiber to fabric assem- Fiber Materials Incorporated. Recent progress blies produced by chemical, thermal or in the automation of the nonwoven 3-D fabric mechanical bonding or a combination of the manufacturing process was made in France by above. Starting with discrete fibers or continu- Aerospatiale (Pastenbaugh, 1988), in Japan by ous filaments (mostly tows), the fibers are Fukuta of the Research Institute for Polymers randomly distributed or preferentially ori- and Textiles (Fukuta et al., 1982 and 1984) and ented by dynamic combing (carding) or more recently by Mohammed (1989). Orthogonal nonwoven (ON) 3-D fabrics are hydrodynamic (waterjet) methods. The Novoltex@structure developed by SEI' fabricated by maintaining one stationary axis as shown in Fig. 18.33 (Geoghegan, 1988)is an either by yarn pre-deposition or a spacer rod example of a mechanically bonded structure which is subsequently retracted and replaced wherein multiple layers of oriented or random by an axial yam. The placement of the planar fiber webs are needled together to create an yarn systems is carried out by inserting the integrated structure which has through yarns orthogonal to the axial yarn system in an
Nonwoven 419
Fig. 18.34 Orthogonal nonwoven by direct method.
alternating manner. In Fig. 18.34, the method. 18.6.2 STRUCTURAL GEOMETRY of direct formation of ON 3-D fabric is shown Structural geomeh.ies resulting from the vari(Stover et al., 1971)* By proper Of ous processing techniques are shown in Fig. yarns prior to planar yarn placement, 18.35: (a) and (b) show the single bundle xyz the 3-D fabrics of various shapes and densities can fabrics in a rectangular and cylindrical shape; be produced. (c) demonstrates the multiple yarn bundle possibilities in the various directions.
Fig. 18.35 Orthogonal nonwoven fabrics.
420 Textile preforming 18.6.3 DESIGN METHODOLOGY
A unit cell geometry for the orthogonal nonwoven 3-D fabric is shown in Fig. 18.36, assuming circular cross-section for yarns in all three directions. The fiber volume fraction for
against d y / d x ratios, assuming a fiber packing fraction of 0.8. For all three levels of dJdx ratios, the fiber volume fraction first decreases with the increase in d y / d xratio, reaches a minimum and then increases. As can be seen in the figure, the maximum fiber volume fraction is about 0.63 at either high or low d y / d x ratios, whereas the minimum fiber volume fraction of about 0.47 is achieved when both d y / d x and d z / d xratios are equal to 1. 18.7 SUMMARY AND CONCLUSIONS
In this chapter, we have first discussed the role and importance of textile preforms in composite design processing and design, followed by classifying them into linear (1-D), planar (2-D) Fig. 18.36 Unit cell for orthogonal nonwoven strucand three-dimensional (3-D) fibrous assemture. blies. The objective of this chapter is to the 3-D ON structure can be shown to have the describe the design methodology of the fiber architecture for representative textile preform following form: structures currently used for composite reinn forcements. After a brief introduction to the Vf = -ICx 2 formation technology of each preform, its fabric structure is shown and the geometry of a dx‘(dy+ dZ)+ d$dX + dZ)+ q d x + dy) (18.23) unit cell is defined. The relationship between (dx + dy)(dx+ dz)(dy+ dZ) the engineering parameters (V,,0) and the key where dx, dy and dZ are diameters of the yarns processing variables (such as preform pattern, in x, y and z directions, respectively and IC is tightness factor 9 and linear density ratio etc.) within the range of achievable geometry is the fiber packing fraction of the yarns. Figure 18.37 plots the fiber volume fraction established from the geometric model. A summary of preform fabrication techniques has been given in Table 18.2. Table 18.3 0.65 gives a summary of the engineering and processing parameters. Ranges of fiber 0.60 orientation angle and fiber volume fraction for 5 each fabric preform commonly used for com0.55 posite reinforcements are also included in Table 18.3. It should be noted that although the achievable range of fiber volume fraction is 0.50 restricted by theoretical fabric geometric limits Minimum fiber volume fraction due to yarn jamming, it is possible that a 0.45 1 . ...I.. ._ ._I higher fiber volume fraction can be achieved 10.’ loo 10’ lo2 lo3 lo4 in reality because of the compressible nature of the preforms. A composite having a higher Fig. 18.37 Process window of fiber volume fraction fiber volume fraction can be made simply by squeezing the preform to a smaller mold for orthogonal nonwoven fabrics. ,
,
,
,
Summay and conclusions 421 Table 18.3 Engineering and processing parameters for textile preforms
Preform Linear assembly Roving Yam
Woven 2-D Biaxial 2-D Triaxial 3-D Woven
Non-woven
2-D Non-woven 3-D Orthogonal
Knit
Fiber orientation, 0 (") 0 - yarn surface helix angle 0=0 0=5-10
0.6 0.7
0, - yarn orientation in fabric plane Oc - yarn crimp angle Of = 0/90, Oc = 30 60 Of= 0/90/+30-60, Oc = 30 60 O6 = 0/90, 0- = 30 60
-
Vf
Processing parameter bundle
- 0.8 - 0.9
Bundle tension, transverse compression, fiber diameter, number of fibers, twist level
- 0.5 - 0.5 - 0.6
Ox - fiber/yam orientation along x axis 0, - fiber/yam orientation along y axis O2- fiber/yarn orientation along z axis Oxy - fiber distribution on fabric plane
Oxy = uniform distribution,
Or,O,O,
= 0
-
Fiber packing in yarn, fabric tightness factor, yam linear density ratios, pitch count, stitch pattern
- 0.7 - 0.6
Fiber packing in yam, fabric tightness factor, braid diameter, pitch length, braiding pattern, carrier number
Os - stitch yarn orientation
Braid
Os = 30 Os = 30
- 60 - 60,
Oi = 0/90/+30-60
0.2 0.3 0.3 0.6
-
0 -braiding angle 2-D Braid 3-D Braid
-
0 = 10 80 0 = 10 45
(2-D non-woven) fiber packing in fabric, fiber distribution (3-D orthogonal) fiber packing in yam, yam cross section, yarn linear density ratios
0.2 0.4 0.4 - 0.6
Oi - insertion yam orientation 2-D Weft knit 3-D MWK
Fiber packing in yarn, fabric tightness factor, yam linear density ratios, pitch count, weaving pattern
0.5 0.4
during the process of matrix addition; how- Assuming a tightness factor 7 of 0.573, possiever, a composite with a fiber volume fraction ble braiding angles range from 0 to 40". Young's moduli and Poisson's ratios of fiber higher than theoretical maximum will have a certain degree of fiber crimp and its fiber ori- and matrix are given as E, = 33.5 Msi, Em = 1.3 Msi, 2rf = 0.3 and urn= 0.11. The elastic conentation will also be distorted. The geometric models of textile preforms stants of the carbon-carbon composite was presented in this chapter provide a quantitative obtained from the Fabric Geometric Model communication link between the preform (FGM) (KO et al., 1987).Figure 18.38 shows the manufacturer, composite processors and prod- composite stiffness in different directions uct design engineers. By reducing fiber within the working window of fiber volume architecture and textile preforming processes fraction and fiber orientation. As can be seen, into engineering and processing parameters Young's modulus, Edav in the axial direction Vf, 8 and 17, rational composite design proce- decreases and in-plane shear modulus, Gh- lane, dures and process control guides can be increases with the increase in braiding angfe 8. established. For example, the mechanistic Young's moduli in both hoop and radial direcdesign of a composite product can be demon- tions, Ehmpand Eradial,have the same value at strated using a tubular 3-D braided zero braiding angle, but depart and both carbon-carbon composite as an example. increase as the braiding angle become higher.
422 Textile preforming 0.006 Braid Axis
Pressure drop: 60 psd
s2 -r?
e.-x 0.004c ._
L2
0.002Limiting fiber architecture
6-
2 0.000 0 _1
3
6
9
12
15
Fiber diameter (pm) 0, 0
,
5
.
.
, . , . , , , . , . 10 15 20 25 30 35 40 45 50 55 60 ~
Braid angle, 8 (")
Fig. 18.38 Stiffness properties of 3-D braided carbon-carbon composite. Fiber packing and tightness factor are assumed as IC = 0.785,~= 0.573; Young's moduli and Poisson's ratios of fiber and matrix are given as E , = 33.5 Msi, Em = 1.3Msi, z),= 0.3, urn= 011.
The other example of the application of the fiber architecture models to the composite processing is to predict the permeability of fabric preforms. As suggested by the well known Kozeny-Carman equation, there are two major geometric parameters which greatly affect permeability of fibrous materials, i.e. porosity of fabric preforms E and characteristic dimension of fibers @D,where @ is the shape factor and D is the fiber diameter. Other parameters which also affect the permeability are flow properties, pressure drop and part thickness. These parameters have been shown to be independent of preform fiber architecture. From geometric analysis, one can construct the V f 4 relationship and determine their dependence on the process parameters. The fabric porosity can easily be calculated from V , ( E = 1 - V,), whereas the shape factor of fibers @ is related to the fiber orientation 8 and the flow direction in composite processing. From our experiments, we have observed that @ = 1.5 when most of the fibers are aligned parallel to the flow direction and @ = 0.75 when the fibers are perpendicular to the flow direction. The effect of the fiber volume fraction and fiber diameter on the permeability of air flow
Fig. 18.39 Effect of fiber volume fraction and diameter on preform permeability.
can be quantified using the Kozeny-Carman equation, as shown in Fig. 18.39, noting that the permeability is in the unit of mass flow rate per length of preform. In this example, the shape factor @ is assumed to be 0.75 at a pressure drop across the preform of 60 psi with a fabric thickness of 0.5 in. REFERENCES Brookstein, D.S. 1990. Interlocked Fiber Architecture: Braided and Woven. Proc. 35th Intern. SAMPE Symposium, Society for the Advancement of Material and Process Engineering, Vol35, pp. 746-756. Brown, R.T., Patterson, G.A. and Carper, D.M. 1988. Performance of 3-D Braided Composite Structures. Proceedings of the Third Structural Textile Symposium, Drexel University, Philadelphia, PA. Chou, T.W. and KO,F.K., eds. 1989. Textile Structural Composites. New York Elsevier. Dow, N.F. and Tranfield, G. 1970. Preliminary Investigations of Feasibility of Weaving Triaxial Fabrics (Dow Weave). Textile Research Journal, pp. 986-998. Dow, N.F. 1985. Woven Fabric Reinforced Composites for Automotive Applications. Technical Final Report, NSF Grant No. DMR8212867, MSC TFR 1605/8102, December. Dow, N.F. and Ramnath, V. 1987.Analysis of Woven Fabrics for Reinforced Composite Materials. NASA Contract Report 178275. Dow, R.M. 1989. New Concept for Multiple Directional Fabric Formation. Proc. 21st Intern. S A M P E Tech. Conf., September 2.528.
References 423 Du, G.W., Popper, P. and Chou, T.W. 1990. Process Model of Circular Braiding for ComplexShaped Preform Manufacturing. Proc. Symp. on Processing of Polymers and Polymeric Composites, ASME Winter Annual Meeting, Dallas, Texas, NOV.25-31. Du, G.W., Popper, P. and Chou, T.W. 1991. Analysis of Textile Preforms for Multi-directional Reinforcement of Composites. 1. Mater. Sci. 26: 3438-3448. Du, G.W. and KO, F.K. 1992. Analysis of Multiaxial Warp Knitted Preforms for Composite Reinforcement. Proc. Textile Composites in Building Construction 2nd Inter. Symp., Lyon, France, June 23-25. Du, G.W. and KO,EK. 1993a. Unit Cell Geometry of 3-D Braided Structures. J. Rein. Plus. Comp. 12 (2): pp. 752-765. Du, G.W. and KO, EK. 1993b. Analysis And Design Of 2-D Braided Preforms For Composite Reinforcement. Proc. ICCM-9, Madrid, Spain, July 12-16. Fukuta, K., Aoki, E. and Nagatsuka, Y. 1984. 3-D Fabrics for Structural Composites. 15th Textile Res. Symp., The Textile Machinery Society of Japan, Osaka, Japan. Fukuta, K., Onooka, R., Aoki, E. and Nagatsuka, Y. 1982. Application of Latticed Structural Composite Materials with Three Dimensional Fabrics to Artificial Bones. Bull. Res. Inst. Polym. Textiles. 131(2)2:151. Geoghegan, P.J. 1988. DuPont Ceramics for Structural Applications - the SEP Noveltex Technology. 3rd Textile Structural Composites Symp., Philadelphia, PA. June 1-2. Goswami, B.G., Martindale, J.G. and Scardino, EL. 1977. Textile Yarns, Technology, Structure and Applications. New York John Wiley and Sons, pp. 273-337. Hearle, J.W.S., Grosberg, P. and Backer S. 1969. Structural Mechanics of Fibers, Yarns and Fabrics, Vol 1,New York: Wiley-Interscience. Kaswell, E.R., ed. 1963. Wellington-Sears Handbook of Industrial Textiles. New York: Wellington-Sears. KO, F.K., Bruner, J., Pastore, A. and Scardino, E 1980. Development of Multi-Bar Weft Insertion Warp Knit Fabric for Industrial Applications. ASME Paper No 90-TEXT-7, October. KO, F.K., Krauland, K. and Scardino, F. 1982. Weft Insertion Warp Knit for Hybrid Composites. Proc. 4th Intern. Conf. Composites. KO, F.K., Fang, P. and Pastore, C. 1985. Multilayer Multidirectional Warp Knit Fabrics for Industrial
Applications. J. Industrial Fabrics 4(2). KO, F.K., Pastore, C.M., Yang, J.M. and Chou, T.W. 1986. Structure and Properties of Multidirectional Warp Knit Fabric Reinforced Composites. In Composites '86: Recent Advances in Japan and the United States, eds. Kawata, K., Umekawa, S. and Kobayashi, A. Proceedings, Japan-U.S. CCM-111, Tokyo, pp. 21-28. KO, F.K., Pastore, C.M., Lei, Charles and Whyte, D.W. 1987. A Fabric Geometry Model for 3-D Braid Reinforced Composites. Intern. S A M P E Metals Conference: Competitive Advancements in Metals/ Metals Processing. KO, F.K. 1988. Braiding, Engineering Materials Handbook, Vol 1, Composites, ed. Reinhart, T.J. Metal Park, OH: AMS International, pp. 519-528. KO, F.K., Whyte, D.W. and Pastore, C.M. 1988a. Control of Fiber Architecture for Tough NetShaped Structural Composites. MiCon '86: Optimization of Processing, Properties and Service Performance Through Microstructural Control, ASTM STP 979, eds. Bramfitt, B.L., Benn, R.C., Brinkman, C.R. and Vander Voort, G.F. Philadelphia: ASTh4 pp. 290-298. KO,F.K. and Kutz, J. 198813.Multiaxial Warp Knit for Advanced Composites. Proc. 4th Ann. Con$ Adv. Composites, ASM International, pp. 377-384. KO, F.K. 1989. Preform Fiber Architecture for Ceramic Matrix Preforms. Ceramic Bulletin 68 (2): 401414. KO, F.K. and Du, G.W. 1992. Processing and Structures of Textile Preforms for Composites. Proc. Science and Innovation in Polymer Composites Processing, MIT, Cambridge, MA, July 16-17. Krcma, R. 1971. Manual of Nonwovens. Manchester, UK Textile Trade Press. Li, W., Hammad, H. and El-Shiekh, a. 1990. Structural Analysis of 3-D Braided Preforms for Composites, Part I: The Four-Step Preforms. I. Text. Inst. 81:491-514. Lord, P.R. and Mohamed, M.H. 1973. Weaving: Conversion of Yarn to Fabric. Durham, UK: Merrow Technical Library. Loos, A.C., Weidermann, M.H. and Kranbuchi, D.E. 1991. Processing of Advanced Textile Structural Composites by RTM. Proc. 5th Textile Structural Composites Symp., Drexel University, Philadelphia, PA, Dec. 4-6. McCarthy, S. and Kim, Y.R. 1991. Resin Flow Through Fiber Reinforcement During Composite Processing. Proc. 5th Textile
424 Texfilepreforming Structural Composites Symp., Drexel University, Philadelphia, PA, Dec. 4-6. Mohammed, M.H., Zhang, Z. and Dickinson, L. 1989.3-DWeaving of Net Shapes. Proc. Zst Japan Intern. SAMPE Symp., Nov. 28-Dec. 1. Pastenbaugh, J. 1988. Aerospatiale Technology. Proc. 3rd Textile Structural Composites Symp., Drexel University, Philadelphia, PA, June 1-2. Pastore, C.M. and Cai, Y.J. 1990a. Applications of Computer Aided Geometric Modeling for Textile Structural Composites. Proc. 2nd Intern. Conf. Computer Aided Design in Composite Material Technology, Brussels, Belgium, April 25-27. Pastore, C.M. and KO, F.K. 1990b. Modeling of Textile Structural Composites, Part I: Processing-Science Model for ThreeDimensional Braiding. J. Text. Inst. 81: 480-490. Popper, P. and McConnell, R. 1987. A New 3-D Braid for Integrated Parts Manufacturing and Improved Delamination Resistance - The 2-Step Method. 32nd Intern. SAMPE Symp. Exhib., pp. 92-103. Potter, K.D. 1979. The Influence of Accurate Stretch Data for Reinforcements on the Production of Complex Structural Mouldings. Composites, 10, pp. 161-167, IPC Business Press Ltd, July.
Raz, S. 1987. Warp Knitting Production. Heidelberg, Germany: Melliand. Scardino, F.L. 1989. Introduction to Textile Structures. In Textile Structural Composites, eds. Chou, T.W. and KO, F.K. Amsterdam: Elsevier, pp. 1-26. Scardino, EL. and KO, EK. 1981. Triaxial Woven Fabrics. Textile Research Journal 51(2). Smith, B.F. and Block, I. 1982. Textile In Perspective. Englewood Cliff, New Jersey: Prentice-Hall. Spencer, D.J. 1983. Knitting Technology. New York: Pergamon Press. Stover, E.R., Mark, W.C., Marfowitz, I. and Mueller, W. 1971. Preparation of an OmniweaveReinforced Carbon-Carbon Cylinder as a Candidate for Evaluation in the Advanced Heat Shield Screening Program. AFML TR-70-283, Mar. Svedova, J., ed. 1990. Industrial Textiles.Amsterdam: Elsevier. Tampol’skii, Y., Zhigun, I.G. and Polikov, B.a. 1987. Spatially Reinforced Composites. Pennsylvania: Teknomic. (English translation, 1992). Williams, D.J. 1978. New knitting methods offer continuous structures. Advance Composites Engineering, Summer, pp. 12-13.
TABLE ROLLING OF COMPOSITE TUBES
19
John T. K a m e and Jerome S. Berg
19.1 INTRODUCTION
In the field of composites fabrication table rolling is a major technique for utilizing preimpregnated fibrous tapes in flag or pennant form for tubular structures. The individual flags become part of the total wall thickness by rolling the flags around a mandrel. The hard mandrel provides the support during cure and defines the inside dimensions of the tube. Table rolling is utilized to fabricate a variety of products including straight tubes usually under 7.62cm (3in) in diameter and up to 3.66 m (12 ft) long and small diameter tapered tubes such as fishing rods, golf shafts and ski poles. Flags may consist of a wide variety of fibers oriented either longitudinally (along the axis of the tube) or offset at a bias angle, hence the terms ’longitudinal and bias flags’. The resin content (RC) and the fiber areal weight (FAW) of the prepreg define the ply thickness. Since external molds are seldom used for table rolled tubes, a variety of polymer compaction tapes are used to apply an external pressure. These tapes provide the external pressure necessary to debulk and prevent flag unravelling before cure and to provide some heat driven compaction during cure. 19.1.1 FIBERS AND RESIN
Carbon fiber form 234 M P (34 ~ msi) to 620 M P (90 msi), glass ’E’ or 8s’ aramids, polyethylene and boron are some of the common fibers
we,
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
in table rolled tube manufacture. The most common resin used to coat the fibers is the epoxy blend family, which is formulated for specific product purposes. The resin and fiber are combined and advanced slightly to a selected ’tack‘ (stickiness) level. This fiber and resin combination is called ’prepreg’. Prepreg surface tack has an important adhesive quality in table rolling which permits composite flags to adhere to one another or the mandrel without slipping during the table rolling operation. Prepreg is offered by specification of resin type and roll width. FAW from 130-160 g/m2 and RC from 30-36% are common in table rolling. Higher modulus fibers favor a lighter FAW to ease rolling. The epoxy in prepreg is catalyzed, so care must be exercised in following the prepreg vendor’s storage and handling recommendations. Freezer storage can extend the shelf life; therefore, the material generally arrives on a freezer truck. Because air and moisture are detrimental to the surface tack, it is important to cut and consume the flags as soon as possible after opening and unrolling the prepreg. Dry and low tack prepreg can influence and aggravate flag wrinkles and ply slippage, leading to voids and dimensional problems. Address tack with the prepreg supplier to find a suitable resin formulation for the table ~ roller’s manufacturing environment. equipment and A Partial list of tooling suppliers commonly utilized in table rolling is presented in Table 19.1 at the end of this chapter.
426 Table rolling of composite tubes 19.1.2 DESIGN
Figure 19.1 illustrates the relationship of the mandrel, prepreg flags and diameters in table rolling. 19.2 EQUIPMENT
19.2.1 SHEAR/SHEETER
A number of commercially available power shears have hardened steel blades and include automatic feed mechanisms for the material rolls. Safety guards with interlocks are needed to prevent finger and hand injury (Fig. 19.2).
Fig. 19.2 Photograph of prepreg sheeter. (Courtesy of Century Design Incorporated.)
The tooling used in this machine is the 'Steel Rule Die'. 19.2.2 ROLLER PRESS The steel rule die, the prepreg and an impact The roller press is a machine used to press sev- sheet (usually a soft plastic like polyethylene) eral stacked layers of prepreg tape into are passed through the rotating pressure individual patterns such as tapered pennants. wheels of the roller press. After compaction,
' Fig. 19.1 Diagram of tapered mandrel, bias flags and longitudinal flags.
-p
Of
longIt@Ml
ma,
Equipment 427 the impact sheet is removed to expose a stack of ready-to-assemble flags which are now nested between the blades of the die. The roller press and dies act similarly to a kitchen cookie cutter (Fig. 19.3). A sharp knife and straight edge can act as a prototype or for small scale production.
---'-.-'
---
,.-.. ._."
I
-.
.-
Fig. 19.4 Photograph of table rolling press. (Courtesy of Century Design Incorporated.)
Fig. 19.3 Photograph of a roller press including loading and unloading racks. (Courtesy of Century Design Incorporated.)
19.2.3 ROLLING TABLE
Originally, in the 1940s, prepreg flags were hand cut and hand rolled like cigars to produce the tubular structure. The fishing rod industry was probably first to commonly use rolled tubular structures. In the 1960s several devices with a mobile lower platen were developed. After activating the machine, a pivoted upper platen is lowered down upon the mandrel and a linear motion activator in the lower platen rolls the mandrel into the prepreg flag. These machines permit pressure ranges to be established, gaining maximum compaction and increasing the speed of rolling. The pivoted upper platen permits the combinations of parallel tube or tapered (cone-like)tube rolling. If the mandrel is parallel, then the pivot function of the upper platen will be unnecessary. Figure 19.4 illustrates a commercially available rolling table. Current rolling tables include temperature controlled platens and platens with piano key-
like fingers for achieving uniform pressure on tapered parts. Both flat bed and segmented bed versions are covered with canvas. This pad provides sufficient resiliency and friction to permit flags to roll without slipping, yet conform to the mandrel surfaces. A slight dusting of talcum powder can be used to prevent prepreg from sticking. Table rolling provides tighter and a more uniform compaction of plies than hand rolling. 19.2.4 VERTICAL TAPE WRAPPER
A variety of plastic and/or cellophane tapes, 1.27-2.54 cm (051.0 in) in width are used to compact the table rolled plies of prepreg. Machines used to apply these tapes must permit tape tensioning to debulk the product as the tape is applied. Some applications call for multiple passes through the tape wrapper to increase the tape pressure for better compaction. Additional wraps of tape are needed for thicker wall structures. Frequently, two types of tape may be used: a release tape and a secondary compaction tape. Some tapes have a release backing which can allow a single pass of tape. However, these tapes are generally more expensive. Apply the tape as soon as possible after table rolling to prevent the flags from loosening. Figure 19.5 illustrates a vertical tape wrapping machine.
428 Table rolling of composite tubes
Fig. 19.6 Photograph of horizontal tape wrapping machine. (Courtesy of Century Design Incorporated.) 19.2.6 MANDRELPULLER
Fig. 19.5 Photograph of vertical tape wrapping machine. (Courtesy of Century Design
Incorporated.)
19.2.5 HORIZONTAL TAPE WRAPPER
This machine represents an alternate to the vertical tape wrapper. It is used frequently for longer, heavier parts and also for very flexible mandrels such as fishing rods. The mandrel is affixed to a chuck or mechanical coupling which rotates the parts while tape is applied. The rollers provide support for the part while motion is in place. The single or even dual tape feed spools move with the tape carriage and return to restart position (Fig. 19.6). After cure, the wrapping tapes are removed by slitting the tape longitudinally and peeling the tape away from the cured part. Wrapping tapes are then discarded.
Tubular parts which have been cured over a hard mandrel are all subject to mandrel extraction. The mandrel puller generally connects to a bolt on the larger shank end of the mandrel. The end of the composite tube rests against a stationary block shaped to permit passage of the mandrel but blocking the tubular part. Mandrel pullers are generally hydraulic or pneumatic. Hydraulic pullers offer a controlled extraction speed, while pneumatic pullers are faster and useful in high volume environments. Mandrel withdrawal is generally done prior to tape removal. Figure 19.7 shows a pneumatic mandrel puller. The type of mold release used, correct size of the stationary block and the mass plus integrity of the cured part must be carefully evaluated or end crushing of the part will occur. Also, thin walled tubes or tubes with a high degree of longitudinal plies can crack during mandrel extraction. 19.2.7 CURING OVENS
Ovens used for curing the composite tubes can be either electrically or gas heated and of batch or conveyorized design. Temperatures ranging from 121-191°C (250-375°F) are most common for roll forming prepregs. Consult
Materials 429 19.3 TOOLING 19.3.1 MANDRELS
Fig. 19.7 Photograph of pneumatic mandrel puller. (Courtesy of Century Design Incorporated.)
the prepreg supplier for recommendations on appropriate cure profiles. Fine tuning of the cure profile is often needed to optimize particular roll forming operations and specific products. Ovens with thermocouples are useful in determining hot and cold spots, which may indicate oven regulation for uniform temperature control. This ensures a uniform gel within the part. 19.2.8 CENTERLESS SANDER OR GRINDER
The wrapping tapes can leave a series of spiral indentations approximately 0.5 mm (0.002 in) deep in the composite tube surface. If a smooth surface is desired for cosmetic reasons or for geometry requirements the part can be surface sanded or ground. A centerless sander basically removes a user defined controlled amount of surface material. A centerless grinder provides a more accurate finish dimension. Centerless grinders are common for the high precision required for the tip ends of golf shafts in which a tolerance of f 0.5 mm (a.002 in) is not uncommon. In carbon fiber golf shaft manufacture, these surface finishing techniques are also used to tailor the product stiffness by incrementally removing material along the shaft length. This changes the shaft stiffness characteristics.
The mandrels used for table rolling are usually hardened steel, sometimes aluminum or even composite. The mandrels are designed to support the prepreg during rolling and curing and provide the inside dimensions for the part. Recalling that mandrels must be extracted in 'mandrel pulling', some negative taper is beneficial. Mandrel makers are skilled in the art of metallurgy. They can select the materials and heat treatments necessary to create a mandrel resistent to permanent bending. However, in many instances the mandrels can be restraightened if damaged. Hard plated mandrels generally provide a longer life since scratched or dented surfaces will hamper mandrel removal. 19.3.2 STEEL RULE DIES
These dies incorporate multiple blades embedded in a rigid backing (usually marine grade plywood) which cut the material in the roller press. The prepreg tape (up to 20 layers) is cut between the cutting die blades and a polyethylene sheet sandwiched between the rotating press wheels. Dies with one piece blades provide the best and most continuous cuts (Fig. 19.8(a)). Dies which include weldments (as in a triangular shape flag) generally dull faster since the weldments soften the cutting edges (Fig. 19.8(b)) Ramps can be used between the cutting blades in die designs to keep the roller pressure off the blade ends. The die builder can recommend blade height and cutting edge type best suited for the task. 19.4 MATERIALS 19.4.1 MOLD RELEASES
Generally, mold releases for table rolling mandrels consist of two components: a primary mold release which provides a polymer bond
430 Table rolling of composite tubes 19.4.2 FIBERS
a
L\ A
Rubber Dads (Stacking reference)
C blades utting\-
U
The prepreg tapes can be made from longitudinal tows or woven tows of the following fibers: aramid, glass, carbon and boron. All these fibrous composite tapes can be cut into flags and pennants needed for the table rolling process. However, because of the brittle nature of boron prepreg, boron is most often cut for longitudinal flags. Refer to the appropriate chapter for specific properties of these fibers.
A
19.5 TYPICAL PROBLEMS
19.5.1 VOIDS
Cutting blades Ramps
Fig. 19.8 (a) Diagram of steel rule die for rectangular shaped flags. Ramps prevent blade damage. @) Diagram of steel rule die for a triangular shaped flags. Ramps prevent blade damage. Weldments can cause blade dulling.
Voids are caused by entrapped air which is not evacuated before resin gelation. The presence of voids reduces the strength bearing capabilities of the part, creates stress risers and can contribute to surface finishing and cosmetic problems. Voids are first minimized by working with the prepreg supplier to assure a high quality material with uniform resin content and good 'wet-out' of the fibers. Also, the material suppliers (prepreg and wrapping tapes) must play a key role in developing a cure profile for the specific process and products. Voids are increased by flag wrinkles which are indicative of rolling problems. The capability to perform void content checks (ASTM D3171) and photomicrographs of the laminate can be extremely useful to develop and improve tube processing. Laminate photos are also very useful in operator training. Few laminates are completely void free but void contents lower than 1%are possible with table rolling and tape wrap compaction.
to the mandrel surface to prevent adhesion; and a secondary mold release which acts as a slip agent. The secondary mold release is most beneficial in straight or slightly tapered mandrels and is reapplied between subsequent mandrel turns. A primary release can lose its effectiveness after several hundred turns and must be 19.5.2 DRY AND DIFFICULT TO ROLL stripped off and recoated. A variety of quality MATERIAL mold releases are on the market. The fabricator should work with the release supplier to Prepreg dryness (lack of tack) can be due to low develop a coating program for the specific resin content, resin formulation, ambient condiapplication. Silicone based releases should be tions of the manufacturing environment, the avoided if the tube is subjected to subsequent age or out time of the material. Insufficient tack can cause flag movement during assembly, bonding and painting.
Typical problems 431 wrinkles, voids and parts with a poor surface finish. Resin content and formulation can be adjusted to suit the manufacturing environment. Temperature and humidity control are very helpful in maintaining consistent material tack in the manufacturing shop. Avoid leaving cut patterns exposed since moisture in the air greatly affects the material surface tack and sometimes renders it useless. Consuming the material within two days is a good rule to follow.
Warm lay-up and rolling tables can help increase material rolling ability and are generally adjusted for slight material and environmental changes. Off angle plies are difficult to roll adjacent to the mandrel and the difficulty is magnified by the higher modulus fibers. Tack tape is a narrow strip of reinforced adhesive designed to aid the adhesion of bias plies to the mandrel. Also, solvent based 'tack resins' can be applied to the mandrel to ease application of the first ply. Once the first ply is tightly rolled, however, the material tack is sufficient for subsequent flags.
Table 19.1 Table rolling equipment, material and tooling suppliers in USA Equipment
Tooling
Century Design Incorporated.
Mandrels Lynco Grinding Corporation 5950 Clara Street Bell Gardens, CA 90201
3635 Afton Road San Diego, CA 92123 (619)-292-1212
(213)-773-2858
Materials
Prepreg Newport Adhesives and Composites 1822 Reynolds Avenue Irvine, CA 92714 (714)-253-5680
Fiberite 4300 Jackson Street Greenville,TX 75403 (903)-457-8554
Mold release Frekote Products Dexter Adhesives and Structure Division One Dexter Drive Seabrook, NH 03874 (603)-474-5541
Wrapping tapes Flexicon Pacific, Inc. 856 North Elm Suite J Orange, CA 92667 (714)-%33-9820
Toray 5729 Lakeview Drive, NE Kirkland, WA 98083-2548 (206)427-9029
Cytec Engineered Materials, Inc. 1440 North Kraemer Boulevard Anaheim, CA 92806 (714)-666-4349
Chemlease P.O. Box 540083 Orlando, FL 32854-0083 (407)-425-2066
Dunstone Company, Inc. 2104 Crown View Drive Charlotte,NC 28227 (704)-841-1380
Steel rule dies Ontario Die Company of America 2735 20th Street Box 610397 Port Huron, MI 48061-0397 (810)-987-5060
432 Table rolling of composite tubes 19.5.3 PART SLIPPAGE DURING CURE
19.5.4 EXPOSED SURFACE VOIDS
The viscosity of the resin drops as the heat of cure begins. Occasionally, tapered mandrels and the constriction of the wrapping tape during the cure can force a part to slip down the mandrel. Golf shaft design is highly dependent on mandrel reference position for proper stiffness and geometry requirements. Slippage can first be minimized by designing a short semi-parallel section in the mandrel (as in the butt section of the golf shaft). Slippage is also reduced by overwrapping the tapes onto the mandrel at both ends to secure the part. In addition, the cure profile or the mold release can be adjusted to limit slippage.
Exposed surface voids after sanding or grinding are indicative of poor rolling practices, insufficient lamination pressure and questionable material. Exposed surface voids are sometimes referred to as 'fiber pulls', which have a wood grain appearance on parts with longitudinal surface plies. 19.5.6 LONGITUDINAL PLY WAVINESS
Tapered parts with longitudinally oriented fibers are prone to zones with a wavy or 'fiber wash' appearance. The problem is amplified with multiple taper mandrels and very low viscosity prepregs. Cure profile modifications or alternate resins can reduce the tendency of 'fiber wash'.
RESIN TRANSFER MOLDING
20
Lihwa Fong and S.G. Advani
20.1 INTRODUCTION
ready for its removal from the mold when sufResin Transfer Molding (RTM) is a closed ficient green strength is attained. Processes mold process in which matched male and that are based on similar principles include female molds, preplaced with fiber preform, Structural Reaction Injection Molding (SRIM) are clamped to form composite components. and different versions of vacuum assisted Resin mix is transferred into the cavity RTM (Figs. 20.1 and 20.2). RTM offers the promise of producing low through injection ports at a relatively low prescost composite parts with complex structures sure. Injection pressure is normally less than and large near net shapes. Relatively fast cycle 690 kPa (or 1OOpsi). The displaced air is times with good surface definition and allowed to escape through vents to avoid dry appearance are easily achievable. The ability spots. Cure cycle is dependent on part thickto consolidate parts allows the saving of conness, type of resin system and the temperature siderable amount of time over conventional of the mold and resin system. The part cures in lay-up processes. Since RTM is not limited by the mold, normally heated by controller, and is the size of the autoclave or by pressure, new Mixing Head
r -------1 I 1
I
I I I I
I
I
I
I
I I
I
Pumpunit
---------
Fig. 20.1 Schematic of the RTM process. Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Vent Port
-
434 Resin transfer molding ISOCYANATE
POLYOL
HYDRAULICS
MACHME MODE
HEAD
Fig. 20.2 Basic construction of a SRIM machine.
tooling approaches can be utilized to fabricate Advantages are: large, complicated structures. However, the 0 Class A surface: Surface definition is supedevelopment of the RTM process has not fulrior to lay-up. In addition, using matched filled its full potential. For example, the RTM tools for the mold, one can improve the finprocess is yet to be automated in operations ish of all the surfaces. such as preforming, reinforcement loading, 0 Close tolerance: Parts can be made with betdemolding, and trimming. Therefore, RTM ter reproducibility than with layup. can be considered an intermediate volume 0 Design tailorability: Reinforcement and molding process (Krolewski, 1990). combination of reinforcements can be used Several unresolved issues in RTM encounto meet specific properties. tered by composite engineers are in the areas 0 Fast cycles: Production cycles are much of process automation, preforming, tooling, faster than with layup. mold flow analysis and resin chemistry. 0 Filler: Filler systems can be used to reduce During the last decade, rapid advances in cost, improve fire/smoke performance, surRTM technology development have demonface appearance, and crack resistance. strated the potential of the RTM process for 0 Gel coat: One or both mold surfaces can be producing advanced composite parts. The gel-coated to improve surface performance. advantages and associated disadvantages of 0 Good mechanical properties: Mechanical the RTM process are summarized. As the properties of molded parts are comparable development of this process is rapid, some of to other composite fabrication processes. the disadvantages may be overcome by the 0 Good moldability: Large and complex advances made in this technology. shapes can be made efficiently and inexpen-
RTM process 435
0 0
0
sively. In addition, many mold materials can be used. Inserts: Ribs, bosses, cores, inserts and special reinforcement can be added easily. Labor saving: The skill level of operator is less critical. Low tooling cost: Clamping pressure is low compared to other closed mold operations. Low volatile emission: Volatile emissions are low because RTM is a closed mold process. The worker is not exposed to chemical vapors as with the lay-up process.
Disadvantages are: Mold design: The mold design is critical and requires good tools or great skill. Improper gating or venting may result in defects. Mold filling: Control of flow pattern or resin uniformity is difficult. Radii and edges tend to be resin-rich. Properties are equivalent to those with matched-die molding (assuming proper fiber wetout, etc.), but are not generally as good as with vacuum bagging, filament winding or pultrusion. Reinforcement movement during resin injection is sometimes a problem. In the following sections, the resin transfer molding process is discussed in terms of the unit operations involved, to familiarize readers with the basic steps of the RTM process. The discussion covers details such as materials of construction, mold design, preforming, curing, and demolding. Processing issues are mentioned in each individual unit operation. Relevant variations of RTM such as vacuum assisted resin transfer molding and flexible molding tools are summarized. Process physics is described with emphasis placed on the principles that govern the RTM process; these are applied in the use of computer simulations. Through the design tools such as simulation codes for mold filling analysis, engineers are able to predict or diagnose the problems in gating and venting in the
design stage. The usefulness of such design tools is discussed in detail, giving the relevant advantages and disadvantages. 20.2 RTM PROCESS
The RTM process can be viewed as seven unit operations. The general practice and processing issues are described for each unit operation. 20.2.1 FIBER REINFORCEMENT
Selection of the proper reinforcement type should take into consideration loading condition, part geometry (size, thickness), mechanical properties and surface finish. The quantity of parts demanded also determines the selection. The reinforcement normally carries 90% of the load in a composite and provides over 90% of the stiffness. The reinforcement in a composite can be designed to match the strength requirements of the part. The following characteristics should be considered when selecting fiber reinforcements: Volume fraction: ratio of the volume of a given mass of reinforcement to the volume of the same component after molding; Wash resistance: ability of a reinforcement to withstand movement due to fluid motion or solvation of the reinforcement binder by the resin; Wettability: ability of a reinforcement to reach a condition wherein all voids in the reinforcement are filled with resin; Sizing: most fibers are coated with size for better wettability and bonding but the size may influence the cure kinetics during the manufacturing. Most standard reinforcement materials for composites can be used, but fiberglass, carbon and aramid are the most common in RTM. One requirement is that the reinforcement should hold its shape during the injection
436 Resin transfer molding
phase. Therefore, the reinforcements are generally stitched, woven or bonded together. Reinforcement build-ups in certain areas can easily be included. For example, woven roving and fabric can be combined with continuous strand mat and chopped strand mat in applications where higher strengths are required. Hybrid systems composed of high performance reinforcements such as carbon fiber and aramid fiber can also be incorporated in RTM laminates. Surfacing materials called veils can be used in the preforms to hide the imprint of fibers, for improved surface finish. Another application of surfacing veil is to achieve a resin-rich skin to improve corrosion resistance. Stitched fabrics (Fig. 20.3(a))reduce stresses inherent in the woven roving design and lead to higher compressive strengths in the composite. However, other constructions such as 8-HS style of weave (eight-harness satin weave) in Fig. 20.3(b)have been used because of improved wetting characteristics and compressive strength compared to bidirectional woven fabrics. Continuous strand mat is multi-stranded, laid in swirled configuration. The mats nor-
mally have 4-6 wt.% of thermoplastic binder added. They are thermoformable and can therefore be used for highly complex shapes or when the anticipated volume of production makes them economical. Different sizings can be obtained on many reinforcements. Sizings can be tailored to the type of resin system. Sizings are available that are compatible with epoxy, vinyl esters or polyesters. The strength variation with type of sizing can be as much as 20%, so this factor needs to be considered in the choice of reinforcement. 20.2.2 PREFORM
For a flat part, the preform can be as simple as a stack of reinforcements that fit in the mold cavity. As preforms become more versatile, various means of producing preforms are available. Currently cut-and-sew is commonly used to assemble preforms of various shapes for aerospace applications. Other near net shape techniques include braiding, spray-up and thermoforming (Fig. 20.4).
I
I
Fig. 20.3 (a) Stitched fabric; (b) eight-hamess satin weave.
I t
1 1 1 1 1 1 1 1 1
1 1 1 1 1 1 1
I
RTM process 437
blank preparation zone
heating zone
stamping zone
unloadinghrimming zone
heatingcuring resinbinder
1
1 cooling A
t+
Vayum source
I
demolding
1
N
directed fiber preform
Fig. 20.4 (a) Four station thermoforming preformer; @) spray-up process.
M
'
forming screen
438 Resin transfer molding
If prefabricated preforms are not used, then some means must be found to hold the layers of reinforcement together as they are built up on the tool surface. For example, unidirectional reinforcement is subjected to washing (washing is unplanned reinforcement movement due to resin movement) if proper precautions are not taken to prevent it. To improve conformance of fibers, a tacky resin (e.g. epoxy), dissolved in suitable solvent (e.g. acetone), can be used as a spot glue to hold the reinforcement layers together. The tacky resin will be washed out during the resin injection cycle and will not interfere with the cure of the part. Sometimes veil can be used to hold the layers and prevent washing.
Edge definition: The edges of the composite will be resin rich if the preform is not cut to fit closely to the edge of the cavity or inserts; Fiber distribution:Uniformity of fiber content in preforms, without excessive thinning, wrinkles or folds, is important; Permeability:A measure of resin distribution into the cavity. This quantity is also affected by fiber volume fraction.
Prefabricated preforms can be further bonded together, with or without a core, to achieve part consolidation. For structural composites, this eliminates the need for fasteners and adhesives to assemble discrete parts. New thermoformable reinforcement mats can be Advantages are: used for highly complex shapes or when the 0 Fast loading: Preforms allow fast loading of anticipated volume of production makes them the mold. economical (Carvalho, 1991). 0 Precise fiber placement: Preform placement Design of preforms should go hand in hand can be made precisely without misalignwith part design. For example, preform corners ment. This allows high quality, close are sensitive to radii of the shape. Figure 20.5 tolerance composites for advanced applicashows the thickness reduction of preform over tions molded by the RTM process. different radii. The preform thickness does not 0 Net shape preforms: If thermoformable change appreciably compared to those around reinforcements are used, the stamped prethe corner when the radius is made larger than forms have excellent dimensional stability. a critical value. However, if the radius is less 0 No additional tool: For low production volume of the composite, the tool for preforming can be the same as the tool for molding.
The only disadvantage with use of preforms is that there is an additional unit operation. With the obvious advantages, use of preforms is advisable when volume of production allows their economical use. When designing fiber preforms, following issues should be considered: Corners: The fiber in the bent corner of a preform tends to move to the inside of a radius. This can cause channeling of flow that leads to poor mold filling patterns; Drapability:This characterizes the ability of a fabric or reinforcement mat to conform to contours of the tool;
1 .o
WRm Fig. 20.5 Effect of comer radii on preform thickness (Rmis the recommended radius).
XTM process 439 than this critical value, dramatic movement of the fibers to the inside of the radius occurs. As a result, channeling becomes dominant in the mold filling stage and induces irregular flow patterns. The edge of a preform is another source of the race tracking of resin. In order to avoid the channeling effect, the preform edge should be cut to fit the edge of the mold cavity. The task of obtaining a good edge definition is normally difficult because of the bulkiness of the layers and inter-layer movement (sliding, rotating) during the mold closing when prefabricated preforms are not used. Preforming of fabrics over tool geometry other than simple flat type will induce shear deformation in the fiber reinforcements. For a biaxial reinforcement, shearing of the weave
(Fig. 20.6) is necessary to conform to the contours of the tool. This drapability problem, therefore, has a two-fold significance in RTM. Because of the fiber rearrangement, the nonuniformity of fiber distribution should be accounted for in the design of the composite. Fiber volume fraction and orientation are no longer that of the unreformed reinforcement. Further, such preforms exhibit different characteristics to resin flow. Designers should account for this change in determining the location of vent ports relative to an injection port. In practice, to modify the permeability of preforms, various flow inducing media or mechanisms have been suggested. Application of such high porosity to the preform or inclusion of a runner system in mold design can alter the mold filling pattern.
Fig. 20.6 Draping fabric on a mold causes shearing of the weave.
440 Resin transfer molding Low viscosity: High viscosities can cause mold pressures that are too high in both the The resin used in the RTM process forms the mold and the injection unit. Raising the temmatrix in the composite after solidification. perature of the resin system is effective to The solid structure is a result from polymerlower its viscosity, but pot life may be ization. To select the resin system, one must adversely affected. take into account of the rheological change Sufficient pot life: This is the time it takes and resin cure kinetics. The formulation of the resin system the resin system's viscosity to reach a level depends on many factors. For example, the that no longer be comfortably handled by the resin system can be combined with promoters, equipment. fillers, internal mold releases, pigments, etc. Tg point: The glass transition temperature Typical fillers, such as clay or calcium carbonshould be as high as possible. As a rule of ate, may reduce cost. The optimum viscosity thumb, the glass transition temperature should for RTM should be less than 500 cP s. Mixing is be at least 30°C (50"F), and preferably 55°C normally required to form a suspension. (100"F),higher than the service temperature Properties requirements (mechanical, chemical, fire retardancy, etc.) can also affect Toughness: Toughness in a resin system is resin selection; the resin mix can be formu- exhibited by its tensile elongation. If sufficient lated to meet specific needs. Attributes to look damage tolerance is required, the elongation for in resin systems are: should be at least 3%. 20.2.3 RESIN SYSTEM AND INJECTION
e consistent reactivity;
ability to wet out the reinforcement; e rapid cure after gel.
0
The ester-type resin mix is combined with an appropriate catalyst, such as emulsified BPO, MEKP, cumene hydroperoxide, at the mixing head and transferred into the RTM mold. Low profile additives have been developed especially for polyester resins to improve surface appearance. In addition, epoxies, urethanes, vinyl esters, nylon and other hybrid resins are available for RTM. The newer resins may require modifications to the pumping/injection unit to meter and condition the resin mix prior to injection. These new systems offer a range of cost and performance options for the RTM process. Influencing parameters are viscosity, pot life, tensile modulus, glass transition temperature, tensile elongation and moisture absorbance. In considering a new resin system, the choice of the proper resin system for RTM must satisfy the following system criteria. Failure to meet these criteria usually means that the resin system is impractical for RTM.
Young's modulus: This modulus must be over some threshold value or the composite compression strength will be less than the optimum value. A high tensile modulus is required to adequately support the fiber reinforcement and prevent premature buckling. The effect of the resin system on hot-humid performance is important in the composite part. The modulus of a typical resin remains essentially constant until the temperature is close to the ultimate T , when it falls off to zero. Under wet conditions, the strength of the resin usually falls off at the same rate as the modulus because of the effect of absorbed moisture. Absorbed moisture plasticizes the resin matrix and lowers the strength of the composite in non-fiber dominated directions. The amount of moisture absorbed by the resin matrix should be small, normally less than 2%. This limits the amount of mechanical performance degradation at elevated temperatures. One final topic to consider is the injection of the resin system (schematics shown in Figs.
Mold materials 441 20.1 and 20.2). Items to control in the resin mix to assure a consistent, smooth running process include: 0 0 0 0
resin mix temperature; ratio of catalyst or curatiire to resin; resin mix viscosity; amount of air entrained in the resin mix. Presence of air in the mix can lengthen the gel time/induce porosity in the composite and/or affect the mix viscosity.
(a)
Most successful production resin transfer molding operations are now based on the use of resin/catalyst mixing machinery using positive displacement piston-type pumping equipment for accurate control of the resin to catalyst ratio. Back pressure at the mix head may change when a mixed resin is injected into a cavity filled with the fiber reinforcement. Static mixers greatly simplify the process and are easily cleaned at the end of the injection cycle. A static mixer sends the proportioned resin and catalyst through flexible Fig. 20.7 (a) Matched mold with rigid halves; hoses to an injection head employing a (b) matched mold with a flexible mold half. motionless mixer to thoroughly blend the materials together immediately prior to injecGate and vent: This critical part of the mold tion step. design should allow complete wetout with minimal resin wastage. 20.2.4 MOLD
RTM mold design and construction is the most critical factor in successful resin transfer molding. The mold must be constructed so that resin reaches all areas. RTM molds require special considerations compared to other composite tooling. Figure 20.7 shows two possible configurations in RTM processing. The mold must be designed to account for the following factors:
Mold sealing: A perimeter gasket is necessary to keep void content low. Tight sealing is important when vacuum is used. Heatingkooling: A typical RTM cycle consists of a wide range of temperature for initiating the chemical reaction, curing and final demolding. Hence, proper heating/cooling channels need to be designed.
Mold materials: The material of construction 20.3 MOLD MATERIALS dictates life cycle of mold, temperature control The low pressure requirements of RTM allow and press requirement. the use of more types of mold materials than can be used in other composites manufacturCavity design: The RTh4 mold should consoliing. The choice between metal molds and date as many assembly steps as possible. A good polymeric composite molds is chiefly one of design should take advantage of this ability. volume and processing temperatures. High
442 Resin transfer molding
volume and high temperatures dictate metal molds. Steel, the most suitable mold material, provides superior face life. Aluminum is good for construction of prototype molds since the metal is easy to machine, is lightweight and has a reasonably high heat transfer rate, but also galls easily. Cast aluminum and spraymetal tooling are currently available and can be used for higher volume applications. Cast copper alloys are being considered for use in RTM molds due to the potential for increased throughput via heat management and better durability. Composites, for example reinforced polyester and epoxies, are most frequently used for making RTM molds. They can be expected to last for approximately 2000 parts (Isorca, 1992).Higher production volumes may justify the use of higher cost spray-metal or metal tools. In some cases, the mold must be backed up in order to maintain its shape. Conventionally the backup can be done cost-effectively with core material or steel frames to add rigidity to the cross section and to support composite mold faces. The closure of the mold is achieved by mating of the mold surfaces against a perimeter gasket. Therefore, guide pins are usually employed to align the mold halves both laterally and vertically to keep resin from leaking. Advancement in adapting composite tooling to the needs of RTM is underway. For example, lengthening the life of the composite tool face is desirable and effective to maintain quality while keeping costs low. The factors that cause deterioration of the mold face are temperature fatigue and attack by solvents or mold release agents. An electrolytically or vapor deposited nickel shell is a new technique that will extend face life.
steps. Therefore, the mold designer should incorporate this rule in the design of the mold cavity. Instead of joining several substructures or onto a major structure after molding, it is structurally more effective and efficient to incorporate them into the part before fabrication. This can be easily achieved by joining substructure preforms when practical. In production the number of molds or cavities required is determined by needed throughput. This should take into account the cycle time. For small parts, the designer can incorporate several cavities in a mold. High surface quality with excellent dimensional control can be achieved by electroplating the mold face with nickel. The appearance surface of a part is usually placed on the bottom of the mold. Pinholes are more likely to collect on the top surface. Mold preparation is similar to that used for hand lay-up. Anew tool must to be treated with several coats of release agent. New mold materials provide flexibility in mold design for RTM. For example, to demold a part with vertical sides, it is common to allow several percent draft in the vertical dimension. Flexible silicone rubber has been used for RTM molds in the form of a bladder mold half which is capable of being inflated or deflated depending on the process requirement. During mold filling, the flexible mold wall is pressed against the rigid wall by inflating the bladder with a pre-determined pressure. During the injection cycle, the mold can deform to enhance resin flow. Upon completion of mold filling, the flexible tool can be further inflated to consolidate the composite component. Part removal in this case is easy since the flexible half can be deflated. %s technique allows fabrication of complicated parts that are not ordinarily possible to demold.
20.4 MOLD CAVITY DESIGN
20.5 INJECTIONPORT AND VENT DESIGN
One of the most important design rules for RTM parts is to reduce the number of assembly
The injection port allows the resin to be transferred into the mold (Fig. 20.1) and its design
Heating and cooling design 443 may be critical. The location of inlet ports must allow the resin to reach all areas without bypassing part of the reinforcement. Air vents help control internal pressure, bleed out air and provide a visual indication of mold filling. Race tracking, or channeling, in the mold is usually the reason why the resin bypasses areas of the reinforcement. Since the resin will not flow backwards, this tends to create dry patches. The engineering way to ensure complete initial wetout is to gate the mold correctly in the design. This may be difficult even for an experienced mold designer. Use of computer simulations as a design tool has become popular in conventional injection molding. Without an engineering design tool, gates and vents can be put in the mold after molding some trial parts, but many trial runs may be prohibitive in some applications. In the next section, new engineering tools adapted for RTM mold filling will be discussed to overcome the problem. Mold designers have found that RTM molds must be vented to allow the air within the mold to be pushed out by the resin. Gate at the lowest point and vent at the highest point is generally a good design practice. Experienced designers may use symmetry to design the inlet ports and outlet vents to remove entrapped air. Venting ports must be placed to draw the resin through sections of the part that are difficult to wet out. They are best placed at dead ends where the resin would not flow by itself. After the resin has finished bleeding, both injection and venting ports must be sealed off. This allows pressure to build up in the mold, and forces the resin to further wetout other sections of the part. This packing stage allows the part to gel under pressure, decreasing void content in the finished part. 20.6 SEALING THE MOLD
The perimeter gasket seals the edges of the mold to prevent loss of resin and injection pressure. In addition, it is an absolute necessity
when vacuum is used. Sealing the mold to achieve cavity pressure of 690 kPa (100 psi) or higher is necessary if the void content of the part is to be kept low. The only practical way to accomplish this is to use O-rings. Machining the face of the mold to close tolerances is prohibitively expensive. It is also usually impossible to maintain the mold absolutely flat to achieve a metal-tight seal. O-ring design is well established. The slot has to be cut so that the O-ring can deform when the mold is closed and maintain a seal. Either square or round O-ring grooves can be used. The type of O-ring material used depends on the maximum temperature the 0ring will experience during the fabrication cycle and the type of solvent used to clean the mold. Nitrile rubber material can be used satisfactorily up to 120°C (250°F). Over 120°C silicone rubber can be used to temperature approaching 177°C (350°F). If help is needed in sealing around inlet or outlet tubes, tacky sealant can be used. This type of sealant is useful for making an O-ring where grooves do not exist. 20.7 HEATING AND COOLING DESIGN
The mold should have good temperature control. The RTM mold should be able to heat and cool the part during the fabrication cycle. Most resin systems cure faster at elevated temperatures. During demolding, lowering the temperature is sometimes helpful in removing the part. Even molds that are intended for room temperature-cured resins should be well insulated so that environmental conditions do not change the gel times and viscosity of the resin. Some molds are heated or designed to go into ovens to achieve faster cures at higher temperatures. Normally, the mold is heated and cooled using either hot water or oil. The mold is constructed to allow the heating/cooling fluid to flow through channels (Fig. 20.8) in its interior. The fluid is heated and cooled by conventional means, such as a gas-fired heater and heat
444 Resin transfer molding
heating channel
-
////////////// TCO
topmold platen resin flow
hh
V/HB la
=
oT -oOoOo Ko0o 0o 0
Fig. 20.8 Heating/cooling by flow channels in the RTM mold.
exchanger. For larger molds, the heating and cooling times will be longer if the heat transfer area does not increase in proportion to the weight. At some point, the production cycle time becomes limited by the rate at which heat can be added or removed, and becomes independent of the curing characteristics of the resin system. Under development is low thermal inertia technology that allows the tool face to be heated by electric wires buried in the face. The construction of the mold face is such that the heat flows into the mold face and not outward toward the mold support structure. This is accomplished by use of a foam core that insulates the bulk of the mold from the tool face. This novel technology, if successful, will allow a more instantaneous transfer of heat where it will do the most good - at the mold face. 20.8 MOLD FILLING
Resin injection is to pump the base resin system to a mixing head through either a single or two pot system. Impingement mixing of the components occurs in the mixing head. The catalyzed mix is then pumped through a static mixer which completes the mixing of the two components. The injection nozzle is attached to the injection port on the mold and the resin system is injected into the mold to pack the mold to a predetermined pressure. When the
mold is filled, the pumping system is shut off and immediately flushed, and the part is allowed to cure. Successful configurations demonstrated in the industry show a common factor: that is, the flow of resin is symmetrical about the vent ports, in a manner such that the volume of air left in the reinforcement decreases. This compression effect helps sweep the remaining air out of the part. When the flow path is arranged in such a way that the resin flows into a configuration with increasing volume, there is a tendency to bypass part of the reinforcement. This situation can happen when core material is used. For example, when there is reinforcement on either side of a core, it is possible that slight misalignment in the core thickness will cause dry spots in the part. To overcome this problem, the resin must be introduced on either side of the core simultaneously. Holes may be drilled through the core to allow the resin system to flow to the other side. When this is done, the core floats on the wet reinforcement and equalizes itself. When the injection pressure is too high or reinforcement tends to move in the mold, the following remedies must be considered:
Multiple gates: partition the mold along the flow path such that travel distance for resin is reduced.
Curing 445 Runner system: allows the delivery of resin to various parts of the reinforcement quickly without using high injection pressure. Flexible mold wall: allows the deformation of the bladder wall to facilitate mold filling. There are several techniques to modify the flow patterns. Application of high porosity media on the preform or inclusion of a runner system in mold design can alter the mold filling pattern. This is helpful in reducing injection pressure or displacing air. All resin movement must be accomplished within the time allowed before the onset of gelation. Additionally, the resin injection process should not cause movement of the reinforcement and should be done at low pressure so that the mold will maintain its shape without requiring massive backing. Vacuum may be used to facilitate filling the mold and simultaneously assist in removing air from the laminate. This requires good mold sealing and the use of a vacuum pump. Vacuum up to 740-760 mm Hg (29-30 in Hg) has been reported in assisting RTM mold fill-
ing (Mosher, 1995). Note that the tooling must be large enough to accommodate the perimeter gasket, air vents, injection ports and guide pins. 20.9 CURING
To convert a resin system into useful products it must be cured or cross-linked by chemical reaction into a three dimensional network. The reaction usually involves either a step growth polymerization, a chain growth polymerization, or a combination of both. The accompanying rheological change in the process is shown in Fig. 20.9 (Macosko, 1989). The curing step constitutes a major portion of a typical RTM cycle. During curing, rheological property changes of the resin system and heat transfer between the mold wall and the resin dictate the cure cycle. Simultaneously, modulus and strength begin to build up at a rate depending on the type of resin and catalyst used and the chemical kinetics of the resin system. Curing can continue after the part is demolded.
Matrix
Time
Fig. 20.9 Rheological change during the curing process. (Reproduced from Macosko.)
446 Resin transfer molding Cure cycle is dependent on part thickness, If the adhesion to the mold face is too strong, the ratio of catalyst or curing agent to resin even exceeding the strength of the composite, and the temperature of the mold and the resin it can be reduced by spraying release agents, system. In some cases, the part is removed normally fatty ester soaps or waxes, on the from the mold immediately after gel occurs. mold surface. After the two mold halves separate, the part The part must develop sufficient green can be removed from the cavity. Part removal strength for handling prior to its removal from methods range from the use of plastic/ the mold. Green strength is the strength a comwooden wedges and rubber mallets to the use posite exhibits after the resin gels, but prior to of knock out pins. A mold designed for low complete cure. Gel time is the interval of time throughput with hand operated clamps probetween introduction of catalyst or curing ducing a relatively simple, lightweight part agent to a thermosetting resin and the formawould most likely be removed using a wedge tion of a gel. Typical gel times range from several minutes to about an hour depending and mallet. Sophisticated hydraulic ejection systems can be used for high volume, complex on the factors mentioned above. The glass transition temperature, Tg,for an or heavy parts. To be pushed out, the part RTM resin system depends on thermal history. needs enough green strength to survive conFor a given temperature, the Tg increases dra- siderable bending stresses. The most common test for sufficient bendmatically with time until it levels off. As the curing temperature is raised, the T reaches a ing strength is to fold over a corner of the part steady-state value at a faster rate. &e steady- immediately after demolding. If the corner state value for Tg is a function of the curing survives the bending without cracks or a temperature, and usually approaches the cur- crease, the part is accepted. Otherwise, meaing temperature. However, the limit is sures to improve its green strength include bounded by the degradation temperature of any of the following steps: the resin system. 0 allow the part more time to cure in the mold; 20.10 DEMOLDING AND POST PROCESSING 0 increase the mold temperature; 0 modify or change the resin system, e.g. The minimum the curing step must accomincrease the catalyst level. plish is to develop sufficient green strength so that the part can be removed from the mold. There is often excess resin at the edges of the While cost is an important factor, it is not the part and in the vents. Considerable trimming, only criteria in choosing a method to remove a part of the post processing, is common when part from an RTM mold. For example, part reinforcement is clamped in the parting line. weight and complexity, and throughput are Trimming is required for almost all items important considerations. In many ways, the made by the RTM process. Accurate preform choice of ejection methods parallels the choice placement and precise alignment can reduce of clamping methods. the labor in this step. A few precautions are required to facilitate Postcure, one of the post processing operademolding. Before opening the mold halves, it tions, is used for various reasons. A molding is necessary to release the part from one mold cycle including postcure can increase producsurface. The force required is approximately tion throughput. While postcuring in an oven, that to overcome the adhesive force between the temperature is not restricted to that the mold and the composite. Typically, tears of allowed for the mold materials. Therefore a surface skin or flash, both resin rich, can be higher conversion of reactive groups can be found around the comers or edges of the part. achieved. It can also prevent the reaction
Process physics and use of simulations as a design tool 447 exotherms of a resin system from damaging a composite tool. It is important to hold the part shape during the process of postcuring and cooling to prevent distortion or warpage. 20.11 PROCESS PHYSICS AND USE OF SIMULATIONS AS A DESIGN TOOL
The processing defects addressed in the previous section are often caused by lack of a systematic treatment in RTM part design and process planning. Among the unresolved issues in RTM encountered by composite engineers, those related to the physical processing have developed rapidly during the last decade. The advancement in RTM technology demonstrates the potential of RTM becoming a primary process for producing many composite parts. In this section, the issues in reinforcement preforming, alternative tooling, mold flow analysis, and cure kinetics are revisited. The focus is on the use of models to describe and enhance the understanding of the physical phenomena. The models are built on the experimental evidence and observations, the goal being to reduce the scope of experiments in the engineering applications. Reducing engineering experimentation is achieved by combining three elements: mathematical models, numerical methods and computer software, into a simulation. One example of such software is LIMS (Liquid Injection Molding Simulation) (Advani et al. 1993) which has been developed specifically for mold filling of complex structures in RTM and can be used also as a design tool for manufacturing of complex structural parts as shown in Fig. 20.10. The topics will be presented in the order found in the unit operations of RTM. Draping of reinforcement plays the role of distributing fibers in a way that depends on the tool geometry. Simulation of reinforcement draping allows an engineer to estimate the fiber content distribution. This distribution can change the volume fraction as well as the orientation
Fig. 20.10 Complex structure manufactured by RTM.
in the molded part and therefore is of extreme importance. Tooling and mold construction are critical factors in successful RTM. By considering several alternative configurations, both the injection pressure and the filling time can be reduced. These alternative designs are valuable as the injection pressure tends to rise rapidly when inhomogeneous fiber distributions are present as a result of preforming. 20.11.1 PHYSICS GOVERNING RTM PROCESSING AND NUMERICAL SIMULATIONS
Darcy’s law for flow through porous media is conventionally used to describe the resin flow in the fiber reinforcement. The generalized form expresses the superficial velocity of resin flow in terms of a factor, which is permeability divided by fluid viscosity, multiplied by pressure gradient. This expression together with the mass conservation in the mold are solved together using various numerical methods. A typical example of this method combines finite element method with control volume method (Bruschke and Advani, 1990a, 1990b, 1991, 1994; Advani, 1994; Young et al., 1991a, 1991b). The solution is moved forward in time after
448 Resin transfer molding
surface. The length of the cell segment can be changed as a result of slippage to accomodate this effect. A dome shaped part will serve as an example of this draping simulation. First, a square bidirectional mat is draped. The workpiece is initially configured so that warp and weft tows are perpendicular to each other. Then draping starts at an arbitrary point on the tool. The initial constraints used in this case study are prescribed along the central tows in both the warp and weft directions. The length of the cell segment is assumed to be constant. In the draped configuration shown in Fig. 20.11, the degree of deformation varies from 20.11.2 PREFORMING cell to cell. The minor angles in the preform For bidirectional mats, woven or stitched, range from 90" to a minimum of 35". The shear draping an arbitrary tool surface depends on also results in fiber volume fraction increase two deformation modes: shear deformation up to 70% for the dome. This information can and inter-yarn slip (Potter, 1979).A mat of this assist a designer in material selection, setup of nature is treated as a net that consists of many processing conditions and part design: a cells (Van West, 1990).Therefore, draping over process engineer can use this information to a surface of double curvature requires the net find out where to make necessary cuts in order to map on the surface by changing the internal to accommodate for induced deformation. As angles in each cell. The four sides of a cell are a rule of thumb, formability of preform mat made up of fiber tows. These tows, under the relies on absorption of such deformation by preforming condition, are inextensible. At the reinforcement material. A good material high deformation regions in a reinforcement, can withstand high deformation without slippage may be necessary to drape the tool wrinkle formation.
the pressure field is obtained during the filling process. The pressure solution obtained from the mold flow analysis can be used to position the gate and vent. This lends a design engineer 'infinite' options when facing the task of mold design. The design rules are no longer restricted to the rule of symmetry used by experienced designers to position the inlet and outlet ports. Instead, a composites engineer would be able to optimize the overall design based on criteria such as minimizing the injection pressure.
Fig. 20.11 Draped dome.
Process physics and use of simulations as a design tool 449 20.11.3 ALTERNATIVE TOOLING
or the preform is driven by the pressure difference. Therefore, the equation of motion is a One benefit of this process is that it can con- function of the bladder as well as the reinsolidate several complex three dimensional forcement material (Lucey, 1992). On the parts into one molded piece. The key to preform part, the compressibility of the reinaccomplish this is tool design. From the design forcement in the thickness direction plays a point of view, a flexible mold wall is very major role. On the bladder part, factors such as desirable to mold certain parts with difficult- inertia, damping, and rigidity of the elasor impossible-to-demold geometry. While a tomeric material can also be included when hard tool makes clamping and demolding dif- they are significant. ficult, the flexible mold provides a convenient From lateral compression tests, the alternative for mold design of these types of load-deflection curve of the fiber reinforceparts. Figure 20.10 shows an example of possi- ment material behaves like a nonlinear spring. ble features which may be molded using this The elastic constant of the preform depends on concept. In this part, one can easily see the its state. Preform permeability is a function of small draft angle and the stiffeners which can fiber architecture and porosity. Since the make demolding difficult. Moreover, the porosity of the preform changes with the beads and the flanged opening in the bulk- thickness, the permeability can be expressed in head of this frame are features that are terms of the cavity thickness. impossible to mold using conventional rigid Figure 20.12 was obtained from the numerimolds. cal simulation of two cases: one with rigid walls To avoid unconstrained wall movement, and the other with a flexible mold wall. In the the bladder pressure is higher than the injec- case with rigid mold walls, the pressure drops tion pressure. The motion of the flexible wall linearly with respect to the flow distance. This is
0
0.2
0.4
0.6
0.0
1
X
Fig. 20.12 Pressure distribution in the 1-D mold near the end of mold filling for flexible and rigid tool.
450 Resin transfer molding
caused by the constant permeability of the pre- pressure drop and overall filling time which is form inside the mold. The pressure curve for the impossible to attain simultaneously in convencase with a flexible mold wall reflects the fact tional tooling (Fong and Advani, 1995). that the fluid flow in the filled region exhibits a smaller pressure drop. This reduction is benefi20.11.4 GATING, VENTING AND VOID cial to the molded parts as it causes less fiber CAPTURE washout and preform deformation due to the In this section, computer simulations for RTM resin. Figure 20.13 shows the results of computed mold filling are discussed to overcome the gatgap thickness of the 1-D mold with a flexible ing and venting problem. Mold filling mold wall. The straight line shows the thick- simulation is an effective way of positioning ness in a rigid tool. From this distribution, one injection and vent ports. Gating and venting can see that the gap height is a function of time are critical in the mold design because they during the filling process as well as a function determine whether complete wetout is achievof pressure. Near the injection gate, the resin able under normal operating conditions. A gate designed at the lowest part and vent pressure balances the applied pressure from at the highest point is generally a good practhe bladder and increases the gap thickness to tice to allow the air within the mold to be its maximum in the 1-D mold. As a result, the resistance to the incoming flow has reduced pushed out by the resin. Experienced designsignificantly as shown in the previous figure. ers may use symmetry to design the inlet Through the numerical study, the potential ports and outlet vents. However, the picture is of the flexible tool design has been demon- often complicated by the geometry or the strated. It has the advantage of reducing the presence of inserts. The engineering way to
0.005
0.004
0.003
0.002
0.001
0 0
0.2
0.4
0.6
0.8
X Fig. 20.13 Gap thickness variation in the 1-D mold near the end of mold filling.
1
Process physics and use of simulations as a design tool 451 ensure complete wetout initially is to gate the the design tool. The situation would be much mold correctly in the design. more complicated if mold filling is coupled Figure 20.14(a) shows a square plate with with phenomena such as preform deformation two cutouts in the part. The injection port is and channeling in the corners and along the first positioned at the center of the lowest part. edges. The flow fronts corresponding to the gate At the microscopic level, heterogeneities design are indicated by the curved contours. always exist in the preform media. For examContours in this figure indicate different time ple, the fiber tow may have a permeability steps. For example, the contours closer to the several orders lower than that of the intergate represents area that is filled first and the stices. Therefore, micro-voids form when the contours closer to the vent the last filled orientation of the fiber tow does not allow the region. As a result of colliding flow fronts in displacement of the air inside the tow. A novel the middle and top portion of the part, the fig- approach in mold filling analysis is reported ure demonstrates the capturing of dry patches by modifying the equation of mass conservaor macro-voids. These voids can degrade the tion to account for the fluid absorbed by the properties of the molded composite signifi- fiber tows (Fong and Advani, 1994). Void cantly. Void capturing is important in the entrapment inside tows is found to be depenprocess simulation to avoid formation of such dent on the microstructure, the vent pressure, defects. Figure 20.14(b) shows an alternative and the ratio of the difference in the permedesign that eliminates venting in the middle of ability of the tows and the permeability of the the part. As a result of injection in the corner, preform (Pillai and Advani, 1994; the vent port has to be positioned differently. Ranganathan et al., 1994). This demonstrates the power and simplicity of
Vent Port
4
I Injection port
(b)
I
Injection Port
Fig. 20.14 Design of injection ports: (a) central injection; (b) corner injection.
452 Resin transfer molding 20.11.5 SENSOR CONTROLLED INJECTION
Sensor controlled injection is multiple injection in an 'intelligent' way without involving a complex control algorithm. It requires placement of gates along the flow path at a number of locations. The injection gate is also a sensor capable of detecting the arrival of resin. These gates are then activated or deactivated in the order of first on, first off, and, therefore, allow the mold to fill in a series of steps. For example, Fig. 20.15 is a simple mold which has four injection gates. To help visualize the concept effectively, a 1-D mold is used. The T-column represents different time stages in the filling process. In this example, only one gate is allowed to open at a time. As the injection starts at T1, the first gate is open and the remaining three are closed. As the flow front progresses through the mold, it hits the second gate location at time T2. The injection unit shuts off the resin to gate one and opens the second injection gate. Instead of having the resin flow through the whole length in the mold, the length is divided up into a number of intervals. Therefore, the overall flow resistance decreases as the effort required is for the resin to flow from one gate to the next closest gate in the flow path.
Figure 20.16 shows the pressure calculation from the mold filling simulation. The pressure drops linearly in the one dimensional flow. As the flow front progresses from the inlet toward the vent under a constant flow rate boundary condition at the inlet, the pressure build-up looks like the schematic shown in the lower left figure. For the multiple gate with sensor controlled injection, the pressure at the first gate increases up to a limit when the flow front hits the next sensor. When the next gate is open, the previous gate is shut off. So the pressure build-up is only limited by the length of the interval. Therefore, the maximum pressure seen in the mold is only a fraction of the pressure compared to the lower left figure.
n q--\n E
a
Flow Length
Flow Length
CONTROL SCHEME T1 T2 T3 T4
ON OFF OFF OFF
OFF ON OFF OFF
/
OFF OFF ON OFF
OFF OFF OFF ON
Mold
Fig. 20.16 Representation of pressure during mold filling.
Table 20.1 shows the results from two sets of computer simulations. For either case, only one gate is open at any time during the mold filling stage. The first column uses constant flow rate and the second column uses constant Table 20.1 Comparison of processing parameters
Gate with flow front sensor
PP
Fig. 20.15 Control scheme for a 1-D model.
Single-gate injection Sensor-controlled injection
Slnglegnfr
100% 48%
s2nglegate
100% 36%
Process physics and use of simulations as a design tool 453 pressure. If only one injection gate is used, the pressure under the constant flow rate boundary condition will reach a maximum. Compared to the sensor controlled injection with four gates, the pressure at the gate is only 48% of the pressure reached by the single gate injection. In terms of filling time, the two molds are subject to a constant pressure boundary condition. Results show that the mold filling for the single gate injection takes almost three times that for the sensor-controlled injection. An example is shown in Fig. 20.17, which elaborates on how one can utilize a sensor to eliminate a dry spot during molding. In Fig. 20.17(a), where no sensors are implemented and the injection gate is at the location as shown, a dry spot will appear in the middle of the part. However, an extra gate in the middle as shown in Fig. 20.17(b), if triggered at the point the fluid reaches the midframe, can prevent this void, as indicated by Fig. 20.17(c). This feature is incorporated in a numerical simulation such as LIMS and can be systematically studied for a given geometry to decide the best strategy when in situ sensing capabilities are incorporated in the fabrication phase (Liu et al. (1995)).
t Vent
20.11.6 MOLD FILLING WITH RESIN DELIVERY SYSTEM
Conventionally, an injection port serves as a point source where fluid is pumped. The drawback of a point source is that the pressure value tends to rise rapidly to an extent that could be detrimental to the preform. By extending the point source into other forms proves to be effective in reducing the pressure build-up. To implement this concept, one can use multiple point sources as discussed previously. A line source has been popular in vacuum assisted RTM because of its ability to fill the mold using 1 atm of pressure. A line source may be modified to serve as a runner by allowing more fiberfree space in this delivery system. This is the channeling effect, now used to advantage in mold filling. Further extending the fluid source may possibly yield a 'plane' source. The actual implementation of a plane source may include a high-porosity layer in the stack-up of the reinforcement mats. The layer can possess a permeability several orders higher than that of the fiber preform. The result of this is a three-dimensional mold flow with fluid propagating rapidly through the spreading plane or surface first
1Vent
Fig. 20.17 Use of sensors to eliminatedry spots: (a) no sensor; (b) extra gate sensor; (c) void prevented.
454 Resin transfer molding followed by percolation of the resin through the thickness of the preform. For three-dimensional flow, venting the mold may become less intuitive. In practice, vacuum assistance can provide part of the solution. 20.12 CONCLUSIONS
Resin transfer molding is a practical process for much of the composite industry. The quality of RTM molded parts can equal that by conventional autoclave processes and its economic advantages are obvious. Although the underlying principles of RTM appear at first to be simple, this is often not the case. The challenge for RTM is to bring together the disciplines of preforming, mold design and process development with existing fibers and resins. This can be best achieved through an understanding of the physics governing RTM and by current simulation technology. REFERENCES
Advani S.G., Bruschke, M.V. and Parnas, R., 1994, Resin Transfer Molding, in Flow and Rheology in Polymeric Composites Manufacturing (Ed S.G. Advani,) Amsterdam: Elsevier Publishers, Ch 12, pp. 465-516. Advani S.G., Bruschke, M.V. and Liu, B., LIMS 3.0: Liquid Injection Molding Simulation, User Manual, CCM Report, University of Delaware, Newark, DE 19716. Bruschke, M.V. and Advani, S.G., 1994, A numerical approach to model nonisothermal, viscous flow with free surfaces through fibrous media, Intern. J. Num. Methods Fluids, 19,575-603. Bruschke, M.V. and Advani, S.G., 1991, RTM: Filling simulation of complex three-dimensional shelllike structures, SAMPE Quarterly, 23(1), 2-11. Bruschke, M. and Advani, S.G., 1990, A finite element/control volume approach to mold filling in anisotropic porous media, Polym. Comp., 11, 398-405. Bruschke, M.V. and Advani, S.G., 1990, Mold filling of generalized newtonian fluids in anisotropic porous media, Transport Phenomena in Material Processing, ASME Trans. HTD 132, 149-158. Carvalho, R.L., Personal communication, Manager
of Application Support Laboratory, Fiber Glass Reinforcements Division, Vetrotex CertainTeed Corporation, 1991. Chou, T.W., 1992, Microstructural Design of Fiber Composites, Cambridge: Cambridge University Press, UK. Isorca Inc., 1992, Introduction to Resin Transfer Molding, Society of Plastics Industry, Composites Institute. Fong, L., J. Xu, and Lee, L,J., 1994, Analysis of thermoformable fiber mat preforming in liquid composite molding: study of deformation modes and reinforcement characterization, Polym. Comp., 15, 134. Fong, L. and Advani, S.G., 1994, The role of drapability of fiber preforms in resin transfer molding, Amer. SOC.Comp., Proc. 9th Tech. Conf., 1246. Fong, L., and Lee, L.J., 1994, Analysis of fiber mat preforming in liquid composite molding, preforming induced effects on mold filling, J. Reinf. Plas. Comp., 13, 637. Fong, L., Varma. R.R. and Advani, S.G., 1994, Use of process models and simulations as design tools in molding polymer and polymer composites, The Pacfic Conference on Rheology and Polymer Processing (PCR’94), Kyoto, Japan. Fong, L. and Advani, S.G., 1994, The role of dual permeability of fiber preforms in mold filling simulation of resin transfer molding, Proc. Zst Intern. Conf. Comp. Engng, New Orleans, LA, 17. Fong, L., Liu, B. and Advani, S.G., 1995, Modeling and simulation of resin transfer molding with flexible mold walls, 50th Ann. Conf., SPI, Comp. Inst., Session 3-A. Krolewski, S. and Busch, J., 1990, The competitive position of selected composites fabrication technologies for automotive applications, Proc. 35th Intern. SAMPE Symp., pp. 1761-1771. Lee, S.M. International Encyclopedia of Composites, 1991, New York: VCH, 1991. Liu, B., Bickerton S. and Advani, S.G., 1994, Modeling and simulation of RTM - venting and void formation, Proc. Intern. Conf. Comp. Engng, p. 17. Lucey, A.D. and Carpenter, P.W., 1992,J. Fluid Mech., 234, 121. Macosko, C.W., 1989, RIM, Fundamentals of Reaction Injection Molding, New York: Karl Hanser Verlag. Mosher, I?, 1995, An introduction to vacuumassisted resin transfer molding (SCRIMP), 50th Techn. Conf., SPI, Comp. Inst., Session 8.
References 455 Pillai K. and Advani, S. G., 1994, The role of dual permeability of fiber preforms in resin transfer molding, Proc. 9th Am. SOC.Comp., p. 17. Potter, K.D., 1979, Composites, lg 161. Ranganathan, S., Wise, G.M., Phelan, F.R., Jr., Parnas, R.S. and Advani, S.G., 1994, A numerical and experimental study of the permeability of fiber preforms, Proc. 10th ASM/ESD Adv. Cornp. Con$, 309. Scheidegger, A.E., 1974, The Physics of Flow through Porous Media, University of Toronto Press. Tucker, C.L., 1989, Fundamentals of Computer Modeling for Polymer Processing, New York: Karl Hanser Verlag.
Vanwest, B.P., Pipes, R.B., Keefe, M. and Advani, S.G., 1991, The draping and consolidation of commingled fabrics, Comp. Manufng, 2, pp. 10-21. Young, W.B., Rupel, K., Han, K., Lee, L.J. and Liou, M.J., 1991a, Polym. Comp., 12, 30. Young, W.B., Han, K., Fong, L., Lee, L.J. and Liou, M.J., 1991b, Flow simulation in molds with preplaced fiber mats, Polyrn. Comp., 12,391.
FILAMENT WINDING
21
Yu.M. Tarnopollskii, S.T. Peters, A.I. Bed’
21.1 INTRODUCTION
A winding operation is the basic fabrication technique for forming load-bearing structural elements made of polymer matrix-based fibrous composites, which have the shapes of bodies of revolution. A semifabricated product (uncured preform) of previously impregnated filaments, strands, tapes and fabrics is wound layer by layer with controlled tension onto the mandrel or previous layers. By varying the angle of filament or tape placement, it is possible to control the reinforcement fiber angles within the same layer and through the thickness of the composite wall. During winding, fiber tension generates pressure between layers of uncured composite; this pressure influences the compaction and void content of the article and contributes to more complete utilization of the strength and stiffness of the reinforcing fibers. If the contact pressure is insufficient for compaction of the material, additional layer-by-layer compaction of a semifabricated product must be employed. The wound article must be converted by chemical and/or thermal means to the finished article. With heat treatment, the usual method, the temperature can be constant or can vary with time. The mandrel defines the internal shape of the article. It is removed after curing if the mandrel is not an element of the structure. The winding process is illustrated in Fig. 21.11,*.
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Filament winding is a natural way to combine two-dimensional reinforcement and, with additional processes and devices, threedimensional reinforcement. Advanced processes, combining filament winding and braiding, allow fabrication of spatially sewn structures. The most important groups of wound articles are: thin-walled shells (their thickness is negligible compared to their radius); compound structures, including three-layered and multi-layered shells with a light foam or honeycomb plastic filler; and thick-walled structural elements. For thin-walled shells, it is most important to optimize the reinforcement configuration. For compound surfaces, contact pressure on the interface is a design parameter. Also, there is the problem of monolithicity for thick-walled structural elements, which is closely related to the problem of control of residual stresses. Most composites cannot sustain a significant internal pressure over time without leaking (weeping) through the otherwise sound composite wall. The use of internal liners made from rubber, plastic or metal can provide a leak tight structure. The semifabricated (uncured) article has extremely low strength in the radial direction. Thus, the ratio of elastic modulii along and across fibers can reach lo3. As additional circuits are wound with tension onto the earlier ’soft’ laminate the previous circuits are compressed with a redistribution of preset tension. The process is typical for all types of filament wound articles with specific processing approaches for each3.Controlling the winding
Introduction 457
q=t(t)
CURING (HEAT BUILDUP, CURING COOLING EXTRA 6OMPACTiAN
TENSIONING
REMOVAL FROM MANDREL
ARTICLE
W
SEMIFABRICATED MATERIAL
INITIAL STRESSES
Fig. 21.1 Stages of winding process.
process may include specifymg fiber tensioning and extra compaction pressure (internal or external), or geometric parameters (configuration of reinforcement lay-up along the length and through the thickness of the article). The reinforcement configuration is determined by the design requirements of the article and can be achieved with high accuracy. Winding, as a way of obtaining preset reinforcement configurations, is restricted in terms of available winding angles. Filament lay-up on geodesic lines on an external surface is easy. (A geodesic line is a line linking two preset points of the surface along a shortest path.) In non-geodesic winding, filament tension can shift the filament
to the geodesic line, against the static friction or viscous resistance of the binder, or mechanical obstacles, such as pins, to keep the filaments on non-geodesic curve. Larger components will result in changes to the system of residual stresses. To calculate, analyze and control the values of residual stresses, it is necessary to find the interrelation between the winding parameters (mainly tensioning) and the properties of the finished product. By controlling filament tensioning and the reinforcement configuration, it is possible to control strength and stiffness of the material, residual stresses and pressures on the mandrel or previous layers.
458 Filament winding 21.2 HISTORY OF WINDING
Historically, the developments in winding technology can be traced back to 1370 BC. There is an ancient Egyptian flask in the British Museum made by winding glass fibers from the melt onto a dissolvable mandrel. Embalmed mummies were wrapped with tapes soaked in resin that were the natural precursors of polyester resins. Rope wrappings were used for reinforcing the bamboo powder rockets in ancient China. Guns with wire wound barrels for reinforcing and proper stress distribution were fabricated in Russia as early as 1869 and in Germany from 1900. In the development of fiber winding technology there have been three important stagesP6:the adoption of the first patint in 19637; the creation of NOL-ring sample4for strength testing in 1955; and the publishing of the first monograph wholly devoted to composite winding technology by Rosato and Grove in 1964R.All these works addressed the development of thin-walled articles rather than the task of converting a semifabricated or uncured material into a monolithic article, which is the primary objective for modern composites technology. The history of analytical development includes the paper by L. Euler on thread equilibrium on a rough surface, papers by P. Laplace and G. Lame on equilibrium of shells loaded by pressure, a paper by A.W. Gadolin on compound cylinders with preliminary tightness and his formulation of the problem on tension wrapping of pressure vessels by wire and papers by A.P. Minakov on thread equilibrium on an arbitrary rough surface. The winding mechanics of an extendible wire have been described by R.V. Southwell. The state of the art of the winding mechanics of composites has been reviewed in detail2.
following stages: winding of a preform; heating of a preform with the mandrel up to temperature of binder polymerization; curing at constant or variable temperature; cooling to below the glass transition temperature and then to room temperature; and removal of the article from the mandrel. As a rule, thermosetting binders are commonly used in winding, although recently progress has been made with advanced thermoplastic binders. Among the reinforcing fibers, the most widely used are fiberglass, carbon, organic (aramid and p~lyethylene)~,~. They can be used in the form of single filaments, rovings, woven and unwoven tapes, strands and woven fabrics. 21.3.2 MATERIALS FOR WINDING
There is a distinction between winding with prepreg, when the reinforcing material is impregnated with a binder by an outside vendor and subsequently partially polymerized to a ' B stage and 'wet' winding, and when the processes of reinforcement impregnation with the binder and the winding operation are performed on the same equipment simultaneously. The rarely used process of 'dry' winding involves winding of the reinforcing material without any binder on the mandrel, followed by vacuum or any other impregnation of the article with the binder and curing. This latter process will see increasing attention since it is a convenient and inexpensive method for providing a stable preform without edges for resin transfer molding. The winding operation of thermoplastics can involve, in the case of in situ consolidation, dilution of the high viscosity of the matrix material by heat or organic solvents. It is physically analogous to the effect of impregnation in 'wet' winding or preliminary heating of a 21.3 TECHNOLOGY OF WINDING prepreg. Although a 'curing' stage is missing in the winding operation of thermoplastics, 21.3.1 STAGES OF PROCESS heat treatment and compaction, as a rule, will The production cycle for most filament lead to significant changes in macromolecular wound composites can be subdivided into the structure. These changes can result in physical
Technology of winding 459 shrinkage analogous to chemical and thermal shrinkage in the curing and cooling of thermosetting plastics. The analysis of winding techniques and general engineering winding theory for thermosetting plastics are surveyed in detail6.
Winding of more complex shapes aggravates the problem of pattern closure. To obtain a continuous layer, it may be necessary to vary the bandwidth during winding. Also, it may be advantageous to vary the fiber volume and thickness by alternate compaction techniques.
21.3.3 GEOMETRY OF WINDING
21.3.4 EQUILIBRIUM CONDITIONS OF FILAMENT ON THE SURFACE
The shape of the article to be wound is determined by the mandrel. The thickness of the article is controlled by the number of circuits wound (taking into account their compressibility during winding and shrinkage during cooling). The reinforcement configuration is determined by controlling the wind angle, which is the angle formed between the tangent to the filament and tangent to the intersection line of the surface to be wound with the plane parallel to the running axis of winding. For a body of revolution it will be the angle between the filament and the wind axis. The formation of a continuous layer with finite thickness and bandwidths, and with proper fiber geometry, requires a coverage solution. If the bandwidth, the number of circuits to close and the winding angle are carefully chosen, the adjacent circuits along the axis will be linked in butt joints, forming a smooth continuous covering. If the bandwidth is greater, the adjacent circuits will form a lap joint resulting in a rough surface. If the band width is smaller, there will be a gap between the adjacent circuits along the axis. This gap can be partly or completely filled up by either fiber or matrix of subsequent layers. Some structures may be amenable to an incomplete winding pattern. Winding of subsequent circuits is feasible either directly on the preceding ones (winding of ribbed and honeycombed structures, Fig. 21.1), or on the rough outer surface formed from several preceding layers. In the second version, the reinforcing fibers undergo additional bending and interweaving and the structure obtained may have more voids and lower fiber volume than in the case of winding with butt joints.
Winding along geodesic lines is easier than along non-geodesic lines. However, to obtain a continuous covering on h e surface it may be necessary to use non-geodesic winding. This presents the problem of permissible deviations from the geodesic lines. The angle of geodesic deviation can be designated by v, i.e. the angle between the main normal v to the curve, along which the filament has been laid, and the normal to the surface n at the same point. The curve on the surface is defined by the radius of curvature Rc and the radius of twist R, and the radius of normal curvature Rn (projection of curvature radius of the curve on normal n to the surface): l / R n = cosv/Rc and radius of geodesic curvature Rg: l / R g = sinv/R, The equilibrium equation for a filament on the surface has the form (Fig. 21.2): ~
~
dN ds
N Rn
+ FT+dQr =
0
+ F,, + p Q , - P = 0
N -+F
+ p Qg = 0
(21.1)
(21.2) (21.3)
RP
where N is tension of filament, s is the arc coordinate on the curve, F represents surface distributed forces, related to a unit length, p is the linear mass density of filament, Q represents mass distributed forces, referred to a mass unit, including dead weight and inertial forces,
460 Filament winding
Fig. 21.2 Condition of filament equilibrium on the surface.
t is
the basis vector directed along the tangent to the curve and P is the normal surface reaction related to a length unit of filament. Equation (21.l) describes the redistribution of tension along the filament length due to friction, viscous resistance from lower-lying circuits, viscous resistance of above-lying circuits and inertial effects. Equation (21.3) represents the force, which tends to shift the filament onto a geodesic line and viscous resistance of both the lower-lying and abovelying circuits. In the simplest case of the friction law I Fg I 5 kP where k is the friction coefficient, from equation (21.2) (at Fn = Q, = 0) and equation (21.3) (at Q = 0) the following relationship can be derive& -Rn =tanv
of pressure. The pressure p on the convex surface comes from tension of the wound material and is unique to filament winding; p = P / b , where b is the width of the material to be wound. For the simple case of circumferential winding of cylinders with a radius R, this pressure is calculated according to the formula:
N P=m
(21.5)
If this pressure is insufficient (if R is high), it is necessary either to increase the fiber tension to a value not exceeding the allowable strength of the fiber band or tows or to employ additional methods. For helical angle winding (at an angle t to the axis) of a cylinder, it follows (21.4) from equation (21.2) and the relationship
Rg
The ultimate permissible angle of geodesic deviation is equal to v Iarctan k. To obtain a monolithic (fully compacted) material with successful chemical processes (matrix crosslinking) in any type of composite processing, some pressure must be maintained through the completion of the curing stage of the matrix. There are several potential sources
Rn = Rc = R / sin'p that
p
=
N sin2 P
-
bR
(21.6)
The greater the angle deviation from the circumferential the less the compaction pressure
Technology of winding 461 generated by the fiber band or tows being wound. There are a variety of methods for creating additional pressure (internal or external, local or general) by means of rollers, power trowels, autoclaves, wrapping of a fiberglass or metallic tape under tension, wrapping of thermally shrinking materials, extensible or squeezing mandrels, rubber bags, etc. Internal pressure can be applied by an expanding elastomer, by pneumatic pressure to an expanding mandrel or by the selection of a mandrel material with high circumferential expansion properties. 21.3.5 REINFORCEMENT CONFIGURATIONS
There are three basic types of reinforcement configurations: circumferential, helical and polar. There are up to six degrees of freedom between the mandrel and the wind eye in typical advanced winding machines today and more motions are possible. Wind eyes can have three additional degrees of freedom (three rotational motions). The combination of several motions may allow more effective or rapid reinforcement placement on a complex surface.
Helical winding The wind eye reciprocates parallel to the axis of the rotating mandrel and the article (Fig. 21.3) and is the most common technique used for tubular structures. By controlling the ratio of rotational and translational speeds, it is possible to control the wind angles of the reinforcement. Three other techniques can be used: one or more wind eyes are stationary, while the mandrel rotates and translates, the wind eye rotates around the translating mandrel, i.e. winding with a stationary whirling arm type winder which is one of the widely used methods of continuous fabrication of pipe; a wind eye rotates round a stationary mandrel and executes a translational motion along its axis. Winding with a whirling arm type winder is a popular method of insulating metallic pipelines. Circumferential winding is a particular case of helical winding with a wind angle of close to 90" (related to the bandwidth and the mandrel diameter).
Circumferential or hoop winding The wind eye is stationary, except for travel along the length of the mandrel at a bandwidth per revolution and the mandrel together with the article rotates about one axis, i.e. for winding of rings, discs, profiled rings and discs (the ultimate case of profiling is the winding through an opening), short pipes, cones and other bodies of revolution. During circumferential winding, the shape of a cross section is decided either by the shape of the mandrel or by localized inserts, or by varying the bandwidth additional hoop layers. Some large and small-scale articles have been fabricated by rotating the wind eye about the stationary mandrel. Circular rings, ellipses, ovals and other shapes can be fabricated by a circumferential winding technique.
I
r
Fig. 21.3 Lathe type winder. Polar winding This type of winding combines several different winding processes. It has the same combination of motions as helical winding, but the shortest axis is the axis of rotation. This technique combines two rotational motions. During simultaneous rotation of the mandrel around its axis and the wind eye around an axis nearly perpendicular to it, a race-track
462 Filament winding pattern is generated (Fig. 21.4). This process enables the fabrication of vessels with different sized polar openings or closed ends. There are a number of alternative techniques. First, the wind eye executes a rotational motion in two planes. It is a very complex technique and is used only for unique articles. Second, the rotational motion of the mandrel is in two planes (Fig. 21.5), sometimes supplemented by a reciprocating motion of a wind eye or turning of its head.
instead of translation motion of a wind eye, there is a reciprocating motion along a curvilinear trajectory. One version of winding with a whirling arm-type winder is when a wind eye rotates and moves along some curvilinear trajectory. Fourthly, there is a group involving two rotational motions executed alternately (a planar-polar winding). Applications of this method include chord winding (Fig. 21.6) of composite flywheels.
3
1’
Fig. 21.4 Racetrack winder.
Fig. 21.6 Scheme of chord winding. Numbers designate the order of chord lay-up : 1-1’, 2-2’, 3-3’ for chords on the right side of mandrel; 1’4, 2’-3 for chords on the reverse side of mandrel; D is a disc mandrel; H is a pendulum mechanism. Fig. 21.5 Polar winding by the mandrel rotating in two planes.
Thirdly, there is a group involving three motions, two of which are rotational in two planes and one is translational. To this group belong, for example, a race-track type of winding, where a wind eye executes a motion along a closed but not circular trajectory. The trajectory shape affects the variation of the winding angle over the surface of the article. Pendulum winding is analogous to lathe winding, but
The fifth group consists of three rotational motions. Kinematics of this process may involve a combination of two rotations around nonperpendicular axes and a combination of two rotations around two perpendicular, but noncrossing axes (winding of a torus). Even more complicated types of winding including combinations of circumferential, helical, polar winding with axial lay-up and/or braiding have been describedzJOJ1. One of the examples of fabricating large-size articles by a combined method is shown in Fig. 21.7.
Technology of winding 463
Fig. 21.7 Combined winding of large-scale articles.
Winding pattern control
explosion of the sporting goods business all over the world. Many golf shaft manufacturIn the simplest winding machines, control of ers have studied or are studying the possibility the ratio of speeds of different motions is done of using filament winding rather than flag by rigid kinematic coupling (usually gears or wrapping for production of carbon/graphite chains). With modem winding machines difreinforced golf shafts. Filament winding offers ferent motions can be regulated individually lower labor costs, because of the possibility of with greater accuracy by computer control. In winding many shafts at once (Fig. 21.9), the the gear and chain machines, looseness and possibility of lower material costs with unlimwear of the mechanical elements have been ited possibilities for varying the axial and the occasionally responsible for poor winding pattorsional stiffness along with the advantage of terns. Winding inaccuracies can result from no seams, such as would be generated by flag combined fiber delivery and winding machine wrapping with unidirectional tape, along the positional errors. These errors can be analyzed shaft’s length. Tennis racquet manufacturers by winding with a controlled material, such as are also considering a filament wound racfine copper wire, over the mandrel surface. quet, again because of the expected lower Typical feedback resolutions on computer conlabor costs and ease of control of stiffness. trolled winding machines range from Along with the new machines much better lWm/bit to 10” m/bit for linear axes and fiber delivery systems have been offered, with from 0.02 to 0.001 degrees/bit for the rotaelectronic fiber tensioners with roll diameter tional axes. The speed sensitivity has been sensing, provisions for setting tension from greatly improved with the newer control systhe main control panel and rewind capability. tems by tracking both velocity and position for The machine manufacturers have develeach axis. This makes it possible to change oped menu-driven, user friendly computer winding speeds during the winding. programs for which the user inputs need only The standard for the industry now is the 6be part geometry, wind angle, band width and axis filament winding machine (Fig. 21.8). number of circuits desired. These have been built in the USA, UK and One of the most helpful innovations introGermany and are marketed worldwide. These duced by the filament winding machine machines, in general, can be programmed by manufacturers has been the ability to work the manufacturer, by third party software supwith the machine control computer programs pliers or by the customer. The newest venue off-line on a common PC within the standard for the machine manufacturers has been the
464 Filament winding I
1
4
Fig. 21.8 6-Axis computer controlled filament winding machine. Machine motion indicated by arrows. (Courtesy of McClean Anderson, Inc.)
operating systems such as DOS or Windows. This allows the design engineer to participate in the selection of machine parameters such as number of circuits to close and the number of crossovers, which influence structural properties. For much of the history of composite fabrication by filament winding, only the large aerospace manufacturers had computer programs to predict slipping or bridging of the wet fiber band on a non-geodesic path. Now these programs are available from third party suppliers and they include 3-D display of the mandrel and the fiber path, the friction at each section and the deviations from geodesic and desired winding angles. They also display the more mundane information, such as laminate thickness, winding time, length of fiber used and laminate weight for any figure of revolution and general nonsymmetric parts (Fig. 21.10). The programming extends to nonaxisymmetric parts also, such as elbows.
21.3.6 MANDRELS
The mandrel, which determines accurate internal geometry for the component, is generally the only major tool. Low-cost mandrel materials such as cardboard or wood can often be used when winding low-cost routine parts. For critical parts requiring close tolerances, expensive mandrels designed for long-term use may be required. For high-temperature cure 315°C (600"F), graphite mandrels with low thermal expansion may be advantageous, however some attention should be paid to the potential difficulties of mandrel removal. Gas containment pressure vessels often require metal liners because composites are porous; these metal liners can also serve as mandrels'*. Mandrels are either removable or nonremovable (remaining as a part of the wound structure). Removable mandrels are classified according to the removal technique as: 0
entirely removed (for example, tubular mandrels with or without taper and with release agent);
Technology of winding 465 ----.,--*----
*. -
.
~
.
-
. .
.
Fig. 21.9 20-Spindle golf shaft filament winding machine. (Courtesy of ENTEC, Salt Lake City, UT, USA.)
0
0
collapsible (the mandrel is disassembled and removed piece by piece); breakable or soluble (plaster, sand or salts).
The selection of a mandrel involves several trade-offs. These include part size and complexity, size of openings, resin system and its curing cycle and the number of components to be fabricated. The basic requirements for a mandrel, whether it will be removed from the part after winding or remain as a part of the structure, are: 0
0
low-melting temperature alloys. Of these materials, the water-soluble sand mandrels and the breakout/washout plasters are most commonly used.
It must be stiff and strong enough to support its own weight and the weight of the applied composite while resisting the fiber tension pressure from winding and curing. It must be dimensionally stable and should have a thermal coefficient of expansion greater than the transverse coefficient of the composite structure.
Components such as rocket motor cases or pressure vessels have closed or reduced-area end openings that require the use of dissolvable or fusible mandrel materials. Some of these materials are: water-soluble sand mandrels, breakout plasters, soluble plaster, soluble or meltable salts, eutectic salts and
Fig. 21.10 Cadwind display of complex mandrel and fiber path. (Courtesy of Material SA, Brussels, Belgium.)
466 Filament winding
Water-soluble mandrels are primarily used in rocket motor cases and pressure vessels where mandrel removal through small openings is desired. The sand/polyvinyl alcohol (PVA) mixture is cast into female molds that have preassembled components, such as insulation, wind axis, lightening tubes and polar bosses. The sand mixture is cured and the two mandrel ends are assembled and bonded. Spider/plaster mandrels are often used to provide a high-tolerance mandrel surface. The plaster is cured, then overwrapped with tetrafluoroethylene tape or other separator film. Following cure, tooling is removed, the plaster is chipped or washed out and the release tape is removed, leaving the desired inside dimensions. Metal-supported plaster is generally used for relatively large parts of 3-6 m. Segmented collapsible mandrels, used for long tubes, are specialized and expensive, but the advantages of their reusability and the continuous winding process renders expensive tooling worthwhile for high-production applications. The main considerations for mandrels are stiffness, hardness, strength, coefficient of thermal expansion and heat resistance. Mandrel removal from the part is equivalent to application of the pressure of opposite sign and new shearing stresses at the interface between the part and the mandrel and must be done with some care. 21.3.7 RESIDUAL STRESSES
The residual stresses caused by the tension of the reinforcement and compaction undergo essential changes during elevated temperature exposure. The thermoelastic stresses as a result of the difference in thermal expansion of the mandrel and composite article and the anisotropy of thermal expansion of composite in a statically indeterminable type of structure are added to the system of residual stresses. The majority of wound articles end up with these stresses. The resulting fields of residual stress in wound articles are comparable in scale to fields of allowable stresses. Taking into
account these fields, their analysis, development and application of the methods of controlling the fields are important aspects in the manufacturing of defect-free structures, especially thick-walled components. 21.4 THIN-WALLED STRUCTURAL
ELEMENTS 21.4.1 TYPICAL ARTICLES
The basic problems in the design and manufacture of thin-walled composite shells are related to the optimization of shape and reinforcement configuration^'^. Fibrous composites show their best performance under tensile load acting strictly along the fibers, while the matrix serves only to distribute the load uniformly among fibers. A rational reinforcement configuration for a thin-walled composite structure is similar to that of a guyed structure working under the same loads. The winding technique is a natural process and is the most widespread production process for thinwalled structural elements. The most widely used elements are cylindrical shells, such as pipelines which operate under a combination of external or internal pressure, bending and axial tension-compression. Cylindrical sections of missiles, cylindrical shells with semi-spherical closed ends used as tanks or pressure vessels, toroidal shells for electrical applications; and conical shells, which include nose fairings of aircraft and rocket motors, are also common. 21.4.2 SHAPE CONFIGURATIONS OF WOUND PARTS
Filament-wound shapes generally include cylindrical, spherical, conical, or dome-end configurations.These bodies of revolution best exploit the advantages of high-speed winding. Spherical shapes contain the maximum possible volume with minimum surface area. Hence this shape is commonly used for pressure vessels. Because the radius of curvature is
Thin-walled structural elements equal in all directions, the best reinforcement pattern is an isotropic combination of wind angles. This pattern is readily obtained by winding a series of great circles, each stepped out from the preceding pattern by the width of one winding band. Cylinders and shafts can be wound on a cylindrical mandrel by providing an adequate pattern reversal length at each end. The cured part can be removed from the mandrel by using a stripping die. Winding patterns consist of hoops or helical patterns as desired to react to design loads. For torsion-type loading, k45" helicals provide the most efficient load path. Thick-wall effects are a major consideration in the design of composite shafting. With flat cylinders, the width is just a fraction of the thickness. These parts can be fabricated using hoop windings alone. A typical application is in flywheels. Thick-wall effects and delaminations are potential problems that must be carefully considered. Fabrication techniques using programmed winding tension, high strain resin systems and low temperature cures help to overcome these problems. The combination of cylinders with domes is typical of rocket motor cases and pressure vessels. The winding patterns and dome shapes must be carefully chosen to prevent fiber slippage and to react adequately against the pressurizing forces on the domes. Ideal, or isotensoid, domes or end closures for filament-wound vessels have a general surface of revolution that requires a numerical solution. The two types of dome contours generally considered are geodesic (constant stress) and polar (planar) ones, each of which requires its own derivation. The only geometric variables affecting the dome shape are the radius of the cylindrical portion of the pressure vessel and the radius to the center of the filament band at the polar opening. There are several winding considerations. Winding patterns must be stable, the dome must be shaped with properly sized openings and there must be balanced stress fields in a cylindrical section.
467
21.4.3 SPECIFIC FABRICATION FEATURES
One of the main problems in obtaining the designed shape for thin-walled composite structures is the problem of warpage. The bending stiffness of structure is proportional to the cube of the thichess, while residual stresses, as a function of thickness, increase more slowly. Therefore, warpage can be a problem with thin-walled articles. The reasons for warpage are : Inhomogeneity of material properties in structures can be ascribed to local imperfections, inaccuratewind angles, inhomogeneity of temperature field during cure and changes in binder composition or volume within the wall. Mandrel removal; pressure and shearing stresses on the interface between the mandrel and article can be distributed nonuniformly as a result of different thermoelastic behavior of the mandrel and composite article, especially for complex shapes. Anisotropy of composite shrinkage and different relaxation rate of components of residual stresses can cause inhomogeneity and anisotropy of relaxation characteristics.
In the production of thin-walled articles it is most important to have a high degree of wind angle fidelity. 21.4.4 SANDWICH AND COMPOUND THINWALLED STRUCTURES
Sandwich structures incorporate a low density filler to increase bending stiffness to increase bending moment of inertia. Foam and honeycomb plastics are most often used as low density fillers. These materials serve not only to increase the separation between the loadbearing layers, but also to act as sound and heat insulation. Sandwich structures are widely used in aviation and construction. Typical design problems are associated with optimization of the thicknesses of load-bearing
468 Filament winding layers and the filler and selection of the reinforcement angles. When winding the outer load-bearing layer the winding tension has to be low enough to avoid damage to the low density filler. Stress distribution during winding depends essentially on the compliance of the mandrel6.In sandwich structures the outer load-bearing layer of the filler serves as the mandrel for the outer skin lay-up. 21.5 THICK-WALLED STRUCTURES
21.5.1 ENGINEERING THEORY OF WINDING
The development of engineering winding mechanics must involve the study, description and sequential consideration of the peculiarities of composite behavior under loading perpendicular to the plane of reinforcement. During the winding process the physicalmechanical properties of the composite undergo great changes. The composite properties remain linear and practically do not change in the reinforcement direction. The compliance in the transverse direction is essentially nonlinear during the stages of winding, heat build-up and curing of the article and can vary by three orders. The possibility of applying a universal rheological model to such a material is beyond the present state of the art. The engineering approach assumes that the material behavior in each of the stages of the process follows a specific rheological law. Passing from one stage of cure to another results in an abrupt change in
material properties. To define these changes, a simplified hypothesis about the progression of the stress state has been developed. Experiments with tensometric (instrumented) mandrels have served as an impetus for the development of an engineering theory (Fig. 21.11). The distinction between the winding process of a semifabricated composite article and that of an isotropic metallic tape was made by examining an 'integral force parameter' - the dependence of the pressure - on the mandrel related to the number of circuits n being wound. A sigruficantpart of the applied pressure is consumed in the deformation process of the lower-lying circuits. By using instrumented mandrels it has been possible to evaluate the variation in winding pressure for every stage of the process. In the curing process, constant pressure on the mandrel is of special interest and makes it possible to develop several variations of the theory of residual stresses, omitting the polymerization stage. It establishesthe dependence of the residual stresses d,and 0 : on the geometry of the article h / R , material anisotropy p2 = E, / E , winding angle p(x, Y), the number of circuits n, winding tension N(x, r ) and parameters of the curing process (q is contact pressure, T is temperature and t is curing time).
Fig.21.11(b) Pressure variation on the mandrel. 0-1
ncr Fig. 21.11(a) Relative pressure (A)on the mandrel in winding. 1is a metallic tape winding; 2 is a composite winding.
is a winding stage; 1-2 is heating after winding; 2-3 is heating prior to curing; 3 4 is polymerization at constant temperature; 4-5 is cooling on the mandrel; N p I X is the ultimate pressure (- - -) on the mandrel in elastic tape winding, R is a radius of the wound article, q,, is a residual pressure on the mandrel after cooling.
Thick-walled structures 469 Residual stresses in thick-walled articles differ essentially from those in thin-walled structures, both in nature and magnitude. In thin-walled articles these stresses can cause the loss of monolithicity and reduction in load-carrying capacity. To analytically describe the force variation in the winding process, the circular models were based on the simplest linear approximation of the nonlinear strain diagram or (Fig. 21.12). The basis of the solution is that the process of continuous tension winding can be modeled by a system of concentric rings, which are first mounted on the mandrel and then on each other under contact pressure equal to the tensioning of the ith layer. The range of thicknesses and operating pressures, were broadened when it became impossible to consider only one linear section of the stress-strain curve of the material. This required model refinement. The modification of this model has played an essential role in the development of technology of winding thick-walled articles, It was established that the total winding pressure on the mandrel had increased nonlinearly, approaching asymptotically an ultimate value 5
p = No arctan sinh nK KR, ~
where K = P(c/R,), Ri is the inner radius of the ring to be wound and c is the band thickness. The critical number of circuits ncrwas established beyond which the winding pressure on the mandrel did not increase. K
Where excessive force conditions of winding appear, a circular area through the thickness of the article appears, where the circuits pass into compression leading to instability in the form of irregular curvatures and buckling of the reinforcement. These defects are common to thick-walled wound articles.
Fig. 21.12 Models of winding stages. (a) a linear circular model (1965); (b) a spirally circular model (1971); (c) approximation of the deformation characteristics of winding materials used for a nonlinear PI,, are coefficients of circular model (1973); PI, stiffness anisotropy of the composite in the process of winding; (d) a model of a growing body, R = R(w),o is a parameter of growth.
PI,,
21.5.2 METHODS FOR CONTROLLING RESIDUAL STRESSES
Programmed winding Programmed winding involves variation of winding tension in accordance with a specified program with the aim of eliminating fiber buckling and compensating for thermoelastic tensile stresses which arise during cooling of the article with the mandrel. Crack formation and loss of monolithicity can occur at this stage. By properly selecting the correct tension program it is possible to transfer the radial stresses or into compression and eliminate the danger of fiber buckling. The range of variation of tensioning N,,, is limited by the strength of the semifabricated article to be wound, the contact pressure 4 = N,,,/R and by the overall dimension of the part to be wound R. New methods have been employed for developing thick-walled wound structures: winding with layer-by-layer curing, introduction of compensating interlayers
470 Filament winding and winding with extra pressure. These methods, along with programmed winding, have facilitated the design of thick-walled articles with wall thickness comparable and sometimes equal to the inner radius. Another means is reinforcing thick-walled articles in the radial direction with a radial reinforcement. This method makes it possible to significantlyincrease shear stiffness and transverse tensile strength. Winding with extra external or internal pressure
This method can be accomplished by layer-bylayer compaction in the winding process, by spinning rollers, by wrapping fiberglass or metallic tape and by extensible (Fig. 21.13) mandrels with programmed compaction pressures to force extra compaction of the article. Compaction of the wound article can also be performed with these same processes during heating and curing. Control of residual stresses can be achieved by varying the magnitude of extra pressure and the time of removal. For a step-by-step compaction technique the following operations (winding of a part of the article, compacting, pressure release) are repeated. In this case, the residual stresses can be altered by varying the intermediate thicknesses of the article and the magnitudes of the applied pressure in each step.
Fig. 21.13 The extensible mandrel (a) and the preform stretched along the axis (b).
cations of programmed winding. The advantages of layer-by-layer winding however result in an significant increase in the duration of the process of manufacturing for thermoset-polymer composites. Therefore, a consideration of the number of individual layers based on a theoretical analysis is very important for process economics. For intermediate curing, both the usual curing methods Winding with layer-by-layer curing and accelerated ones (i.e. radiative or catalytic) The basic idea of this method is similar to step- are appropriate. by-step compaction in that it decreases the compliance of the already wound circuits. This A method of temperature gradients is a result of partial or complete curing of a group of wound layers, followed by winding There are methods for controlling the residual of the next group of layers. As the number of stresses by varying the process parameters in curing steps, into which the process is divided, the heat build-up and curing stages. increases, layer-by-layer winding approaches Thermoelastic behavior of anisotropic bodies some idealized ultimate process of winding a differs greatly from the thermoelasticbehavior ’cured’ material (similar to thermoplastics). of isotropic bodies. The methods of temperaThe layer-by-layerwinding extends the appli- ture gradients must compensate, at least
Applications 471 partly, for the thermoelastic stresses which arise from anisotropy of thermoelastic properties, by artificially developed and controlled temperature gradients. Different modificaand Objectives for using the method Of temperature gradients are possible. Among them are the curing in an inhomogeneous temperature field, the creation of temperature gradients during the cooling process, or modification Of the Polymer matrix through the thickness of the article for the purpose of controlling the gradient Of heat during a curing chemical reaction.
21.6 APPLICATIONS
21.6.1 PRESSURE VESSELS
Most advancesin filamentwinding have been made in the course of development of pressure vessels, particularly rocket motor cases. In these structures the reinforcement is loaded only in tension and winding the arrangement of fibers along the directions of principal stresses. It uses the main advantages of composites, such as high strength and stiffness in the reinforcementdirection, and avoids shear stresses which are dangerous for composites. Rocket motors come in many sizes, from the graphite/epoxy filament wound case A spatial reinforcement for the Space Shuttle to pressure vessels as A number of methods for controlling residual small as 0.152 m in diameter for orbital adjuststresses at the expense of variation of the reinments to spacecraft. Metal-lined pressure forcement configuration have been worked vessels are widely used in aerospace, aircraft, out. By controlling portions of the elastic, marine and commercial applications. The strength and thermo-physical properties Skylab Oxygen Tank shown in Fig. 21.14 is an either by changing their level or generating example of such a structure. The tank, 1.143 m artificial inhomogeneity, it is possible to affect in diameter and 2.286m long, operated at the level and distribution of residual stresses. 31 MPa with a demonstrated life up to 1000 There are methods for the reduction of cycles in temperature service range of anisotropy of composites, such as an extra -53-71°C. radial reinforcement by short needles, winding while braiding a three-dimensional reinforcing framework, hybrid reinforcement and decreasing of fiber volume content, Another method of controlling the residual stresses is to introduce compliant interlayers of rigid foam or other material. The third group of methods creates an artificial inhomogeneity by substituting a homogeneous structure for a compound one which has different elastic and thermophysical properties. This is done by creating a smooth macro-inhomogeneity of properties by varying the bulk reinforcement coefficient, or the ratio of two types of reinforcement in hybrid composites, or by changing the reinforcement angle or chemical composition of matrix through the article. The basic methods for controlling the residual stresses are summarized in Table 21.1. Fig. 21.14 Skylab Oxygen Tank5. (Courtesy of Brunswick Corporation.)
472 Filament winding Table 21.1 Methods of controlling residual stresses in wound articles Techniques
Method
Forced
A programmed winding Pressing and rolling during winding process and after it
Chemico-technological
Curing in regime of spreading of reaction front Curing in nonhomogeneous temperature field Variation of matrix composition through the article Continuous or layer-by-layer curing during a winding
Controllingphysical fields
Controlling of cooling rate Method of temperature gradients Influence of electromagnetic fields at curing
Controlling of reinforcement schemes
Additional radial reinforcement Introduction of thermocompensatinginterlayers Variation of reinforcement coefficient Application of hybrid composites Creation of artificial nonhomogeneity Creation of compound structures Winding with braiding
21.6.2 THIN-WALLED SUPPORTED
STRUCTURES
amount of attenuation. By increasing the pressure of a water-filled test vessel, deep water depths could be simulated for calibration of sonic equipment, while the equipment is suspended just a few feet under water. A 2.4 m diameter acoustic pressure vessel with 0.10 m wall thickness rated for multiple cycle life 34.4 MPa working pressure is shown in Fig. 21.15. Special winding programs were
Thin metal-lined pressure vessels overwrapped with Kevlaf reinforcement for the Space Shuttle and various medium class satellite systems and metal fuel tanks overwrapped by composites for military aircraft are included here. The applications in aircraft structures include entire composite aircraft fuselages as well as many secondary structures. Two prototype fuselage test sections for Beech Starship with different stiffening concepts have been developed, both filament winding concepts used high strength graphite and Shell resin 9400. The first technique used a filament-wound isogrid network, 1.8 m diameter, 2.4 m long geodesically reinforced test section to prove the concept. The second approach used a honeycomb-stiffened structure with graphite epoxy skins and lightweight composite tooling. One of the earliest uses of composites for marine application was in the use of acoustic transmission of radar signals. Fiberglass-reinforced plastic structures have an ability to Fig. 21.15 Eight foot (2.4m) diameter acoustic prestransmit sound waves with a minimum sure vessel5.(Courtesyof Brunswick Corporation.)
Applications required in concert with multiple compaction and cure cycles due to the thickness of the vessel. Composites combining low density, stiffness and high strength with superior sound damping properties are very promising materials for submarines and deep-submergence devices. Filament wound composites offer cost savings in terms of materials and process both for primary and secondary structures on submarines. 21.6.3 NETTED-RIBBED SHELLS
The middle layer of a sandwich composite structures is usually fabricated by an alternative technology. Winding technology can be used not only to manufacture facing layers, but also for middle layers by winding a network of helical ribs. Netted-ribbed composite shells (Fig. 21.16) are widely used in aerospace because of low dead weight, a9 antenna frames, solar batteries, etc.I4. 21.6.4 WINDING WITH BRAIDING
Winding with braidingi0 allows the manufacture of articles of complex shape with sophisticated reinforcement techniques. These structures possess high resistance to bending, torsion, tension, compression and to local
Fig. 21.16 Carbon-epoxy filament wound isogrid space structure.
473
loads. The advantage of this technology has been demonstrated in high resistance to impact loads. The impact failure is localized within several spatial cells of the structure. 21.6.5 THICK-WALLED STRUCTURES LOADED IN TORSION
Progress has been made in the development of thick-walled wound rods loaded in torsion. A characteristic example is shafting for shipd5. The advances in the mechanics and technology of winding of thick-walled largescale elements have substantially solved this problem. It is well known that fibrous composites behave poorly when loaded in torsion. Ways of increasing load-carrying capacity involve increase in shear strength of a ply and the adherence to the optimal law of shear stress distribution over the radius of a rod. Increase in the strength of a ply is accomplished by properly selecting the optimal angle in the plane of fiber lay-up and the number of starts and crossovers during winding. Shear strength in torsion zuLT= 400 MPa has been attained and a shaft was fabricated with wall thickness equal to the inner radiusi6. Thick-walled wound structures can also be manufactured by multidirectional braiding of a system of one tow. The main difficulties of this process are the complexity of preform impregnation and necessity of compacting over the external surface to increase fiber volume and to ensure monolithicity of the article. This promising technology can also be used for braiding over the external surface of unidirectional thick-walled elements, to reduce the danger of delamination due to low transverse tensile strength. Another possibility to increase the torsion strength of thick-walled products is the spatial placement of straight fibers along the surface contours of one-sheet hyperboloids, whose centers are shifted along the longitudinal shaft axis for a specified step. The slope of straight contours of one-sheet hyperboloid can be governed by the wind angle of the applied fibers, which are originally
474 Filament winding parallel to surfaces. The spatial arrangement of reinforcement results not only in the increase of transverse tension and interlaminar shear strength and in the localization of fracture within several spatial 21.6.6 COMPOSITE FLYWHEELS
The ultimate storage energy W" of a rotating body is proportional to tensile strength of the material at"and body volume V , i.e. (W" = k,oW) and is independent of the material density p. The ultimate angular speed of flywheel depends on the material density and consequently, composite flywheels are high-speed structures. Mass energy storage capacity Wmu is proportional to specific strength of material Wum= WU/m = kma"/p where m is flywheel mass and volume energy storage capacity Wvuis
where V , is the overall flywheel dimension. Advanced composites are the most promising materials for inertial energy storing systems because of high specific strength, high tensile strength along reinforcing fibers and comparatively safe modes of failure. The possibility of quick release of all stored energy is one of the advantages of flywheels. The highest mass energy storage capacity is achieved with rim flywheels made by a circumferential winding technique (km -+ %).This is true only for thin rings. Increasing volume energy storage capacity requires growth of the radial thickness of the rim. With increasing radial thickness the problem becomes bidirectional, i.e. not only circumferential but also radial stresses must be considered. The low transverse tension and shear strengths of unidirectional composites limit the ultimate rim thickness. This is why
Fig. 21.17 A full-scale specimen of the hull of underwater vessel of reinforced plastic before test in a pressure chamber19.
References
previously developed methods of increasing load-carrying capacity of thick-walled articles appear to be extremely useful in the design and fabrication of composite flywheels1s.The high effectiveness of methods of increasing transverse tensile strength of composites by improving the matrix-fiber adhesion, i.e. by plasma surface treatment of carbon fibers, may be of greater value for thick wall fabrications. Use of chord winding is another way to increase radial tensile strength and provides a solution to solve the following problems: how to connect the rim, shaft and hub of the flywheel and how to increase shear strength and stiffness. 21.6.7 DEEP SUBMERGENCE DEVICES
Deep submergence composite devices with the depth of immersion up to 2000 m (manned)and up to 6000 m (unmanned)l9 have been built. A hull with 22m length, 8 m diameter and 250 t total mass was entirely manufactured by winding on numerically controlled machines. The full-scalehull (2.4 m diameter and 17t mass) for an underwater vessel made up entirely of fiberglass plastic (metal fraction in the structure less than O.lo/0) withstood an external pressure 350 MPa is shown in Fig. 21.17. Tests methods for wound articles have been described in detail in References 15 and 20. REFERENCES 1. Peters, S.T., Foral, R.F. and Humphrey, W.D. 1987, Filament winding. In International Encyclopedia of Composites, (ed. S.M. Lee) pp. 503-518. New York: VCH. 2. Structural Composites. A Handbook, 1994, (eds. Yu.M. Tarnopol’skii and V.V. Vasiliev) Lancaster: Technomic. 3. Peters, S.T. 1987, Filament winding. In Engineered Materials Handbook, Vol.1, Composites, pp. 503-509. Ohio: Metals Park. 4. Handbook of Composites, 1982, (ed. G. Lubin) New York Van Nostrand Reinhold. 5. Peters, S.T., Humphrey, W.D. and Foral, R.F. 1991, Filament Winding Composite Structure
475
Fabrication, 2nd printing. Covina: SAMPE. 6. Tarnopol’skii, Yu.M. and Beil’, A.J. 1983, Problems of the mechanics of composite winding. In Handbook of Composites, (eds. A. Kelly and Yu.N. Rabotnov), Vol. 4, Fabrication of Composites, (eds. A. Kelly and S.T. Mileiko) pp. 45-108. Amsterdam: North-Holland. 7. Concise Encyclopedia of Composite Materials, 1989, (ed. Anthony Kelly) Oxford: Pergamon Press. 8. Rosato, D. and Grove, C.. 1964. Filament Winding: Its Development, Manufacture, Applications and Design. New York: John Wiley. 9. Tsai, S.W. 1994, Theory of Composites Design. Dayton, Paris and Tokyo: Think Composites. 10. Textile Structural Composites. 1989, (eds. T.W. Chou and F.K. KO) Composite Materials Series, Vol. 3, (series ed. 8. Pipes) Amsterdam: Elsevier. 11. Tsyplakov, O.G. 1974 (Part l), 1975 (Part 2), Scientific Fundamentals of Technology of Fibrous Composite Materials. Perm: Perm Book Publisher. 12. Humphrey, W.D. and Peters, S.T. 1987, Filament winding. In Engineered Materials Handbook, Vol. 2, Engineering Plastics, pp 368-377. Ohio: Metals Park. 13. Obraztsov, I.F., Vasiliev, V.V. and Bunakov, V.A. 1977. Optimal Reinforcement for Composite Shells of Reuolu tion, Moscow: Mashinostroyenie. 14. Bunakov, V.A. and Protasov, V.D. 1992, Netted composite cylindrical shells. In Proc. 1st USSR-USA Symp., Mechanics of Composites, Vo1.2, Composite Structures, (ed. Yu.M. Tarnopol’skii), pp. 89-98. Riga: Zinatne. 15. Wilhelmi, G., Appelman, W. and Loo, F. 1986, Composite shafting for naval propulsion system. Naval Engng., July, 129-136. 16. Tarnopol‘skii, Yu. M. and Zakrzhevskii, A.M., 1994, Thick-walled composite rods loaded in torsion. Mechanics Comp. Mater 30(4) 40-47. 17. Tarnopol’skii, Yu.M., Zhigun, I.G. and Polyakov, V.A. 1992, Spatially Reinforced Composites. Lancaster: Technomic. 18. Portnov. 1989. Composite flywheels. In Handbook of Composites, (eds. A. Kelly and Yu.N. Rabotnov), Vol. 2, Structures and Design, (eds. C.T. Herakovich and Yu.M. Tarnopol’skii) pp. 531-582. Amsterdam: North-Holland. 19. Advanced Materials and Processes. 1993. Scientific Technical and Advertising Article Collection 1, 70-72. St. Petersburg: CRISM Prometey. 20. Tarnopol’skii, Yu.M. and T. Ya., Kincis. 1985. Static Test Methodsfor Composites. New York: Van Nostrand Reinhold.
FIBER PLACEMENT
22
Don 0. Evans
22.1 INTRODUCTION
Fiber placement is a unique process combining the differential tow payout capability of filament winding and the compaction and cut-restart capabilities of automatic tape laying. During the fiber placement process, individual prepreg fibers, called tows, are pulled off spools and fed through a fiber delivery system into a fiber placement head (Fig. 22.1). Here they are collimated into a single fiber band and laminated onto a work surface which can be mounted between a headstock and tailstock.
FiberPkementHead
When starting a course, the tows are restarted and cogpacted onto a surface. As the course is being laid down, the processing head can cut or restart any of the individual tows. This permits the width of the band to be increased or decreased in increments equal to one tow width. Adjusting the band width avoids excessive gaps or overlaps between adjacent courses. At the end of the course, the remaining tows are cut and the head is positioned to the beginning of the next course. During the placement of a course each tow is dispensed at its own speed, allowing each
r
T i Restart Rollers
Collimated Fiber Band
Y
ControlledHeat
Fig. 22.1 Fiber placement head. Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Introduction 477 tow to independently conform to the surface of the part. For example, when the head laminates a curved path, the outer tows of the fiber band pull more length than the inner tows. A rolling compaction device, combined with heat for tack enhancement, laminates the tows onto the lay-up surface. This action of pressing the tows onto the work surface (or a previously laid ply) adheres the tows to the lay-up surface and removes trapped air, minimizing the need for vacuum debulking. Figure 22.2 is a diagram of a Cincinnati Milacron 'Viper' Fiber Placement System (FPS). This system has seven axes of motion and is CNC controlled. The machine consists of three position axes (carriage, tilt, crossfeed), three orientation axes (yaw, pitch, roll) and an axis to rotate the mandrel. All of these axes are necessary to make sure the processing head is
normal to the surface while the machine is laminating tows. The machine also has 24 programmable electronic bidirectional tensioners, which are mounted in a creel. These tensioners provide individual tow payout and maintain a precise tension. The fiber placement head is mounted on the end of the wrist. The head precisely dispenses, cuts, clamps and restarts individual prepreg tows. To increase productivity some machines are equipped with dual mandrel stations (Fig. 22.3). This setup allows two sets of lay-up tools to be placed on the machine, ensuring a constant supply of work for the head. During routine manual operations such as hand laying small plies, inspecting plies, or vacuum debulking on one part set, the head simply shifts to the other mandrel and picks up the program where it left off.
X-axis (arm crossfeed) Y-axis (arm tilt) Z-axis (carriage longitudinal movement) &axis (pawmandrel rotation) i-axis (head yaw) $-axis (head pitch) -axis (head roll) -axis (redirect roller angular position) -axis (tow restart linear position)
Fig. 22.2 Fiber placement system.
478 Fiber placement
Fig. 22.3 Fiber placement machine dual station.
rest of the tape is despooled. This will cause the tape to eventually break. During part fabrication this backing film is removed before the fiber reaches the fiber placement head. When selecting a tow that is to be used by a fiber placement machine, it is important to consider the cross sectional area of the fiber. This cross sectional area will be a major factor in determining the ply thickness of the laminate. This is because the width of the tow is fixed and the thickness of the tow varies depending on the cross sectional area of the fiber and the resin content. The following equations can be used to determine the thickness of a given tow. Since resin content is usually given as a percent of resin by weight, it is first necessary to calculate the resin volume fraction, which is the percent of resin by volume.
22.2 MATERIALS
A fiber placement machine can dispense prepregged fibers that are commonly used by the aerospace industry such as carbon, KevlaP and glass. These fibers need to be impregnated with a resin and formed into tows or slit tape. The width of tow or slit tape used by fiber placement range from 3.2 mm (0.125 in) to 6.4mm (0.250 in) After the tows are impregnated with a resin matrix they are flattened to a desired width and wound onto a 7.6 cm (3 in) diameter by 28 cm (11 in) long core in a helical pattern. A typical length for a 2.3 kg (5 lb) spool of prepregged IM7-12K tow, 3.2 mm (0.125 in) wide, is 3350 m (11 000 ft). Slit tape is fabricated by running a 7.6 cm (3 in) wide tape through a slitter creating smaller widths of slit tape. These slit tapes are then wound onto a number of 7.6 cm (3 in) diameter by 28 cm (11 in) long cores. When the slit tape is wound onto the spool a backing film which is wider than the slit tape must be added. If the backing film is not used, the slit tape cannot be removed from the spool because of stringers, which will occur during the despooling operation. The edge of the slit tape separates and stays on the spool while the
where rvf = resin volume fraction, rwf = percent of resin by weight fraction, rd = resin density (g/mm3), r f = fiber density (g/mm3). Using the resin volume fraction the thickness of fiber can be calculated using the following equation: (22.2) where fh = tow thickness (mm),fa = fiber cross sectional area (mm2),fw= tow width (mm),rvf = resin volume fraction. For example the thickness of a Hercules IM7 fiber impregnated with a Hercules 8551-7 resin containing a 32% resin content by weight and flattened to 3.2 mm (0.125 in) would be 0.14 mm (0.0055 in). To increase or decrease the thickness of the tow, the cross section of the fiber must be changed by adding or subtracting 6K or 12K fiber bundles. With slit tape this is not a problem because the thickness of the tape is determined when it is prepregged and before it is slit to width.
Tooling considerations 479 The tow width of the material is very guided through a fiber delivery system and important in controlling the gap between the head, but high tack is needed when it is being prepregged tows. For example if the fiber compacted onto the surface. placement head is designed to lay down tows Materials that have a low tack can be that are 3.24.38 cm (0.125+0.015in) wide, the despooled with a fiber tension of 0.45 kg (1lb) tows will be compacted onto the surface in or less. These low tensions are achieved 3.2 mm (0.125in) spacings. If the tow is exactly because the resin does not stick to the spool or 3.2 mm (0.125in) wide, there will be no gap the components of the fiber delivery system. between the tows. If the tows are 2.5mm This lower fiber tension is needed while fiber (0.100in) wide, there will be a 0.63 mm placing concave areas. A higher tension will (0.025in) wide gap between the tows. If the cause the fiber to bridge over concave areas. tows are 3.8 mm (0.150 in) wide, there will be Materials with low tack levels also have less a 0.63 mm (0.025in) overlap. tendency to deform or rope while being pulled Figure 22.4 shows a typical width his- through the fiber delivery system. They also togram of a spool of tow material. To develop transfer less resin to the components of the this histogram a spool of tow was run through fiber delivery system and head. This reduces a measuring device, which measured the the number of times that these components width at 1100 locations. The average width is need to be cleaned because of resin build up. 3.20 mm (0.126 in) and the standard deviation Resin build up in the head causes it to malis 0.13 mm (0.005 in). A typical width his- function. togram of a slit tape will show an average The tack of most resins can be reduced by width of 3.18 mm (0.125 in) and a standard lowering their temperature. To accomplish deviation of 0.08 mm (0.003 in). Maintaining this, most fiber placement machines are the quality of tow and slit tape will greatly equipped with an air conditioned creel. The improve the quality of the composite laminate fiber placement head also has some means of cooling the components that come in contact and the reliability of the process. The ideal fiber placement material has no with the tow material. To achieve the desired tack at 21°C (70°F) and high tack from 27°C to tack required to adhere the tow material to the 32°C (80°F to 90°F). Low tack is needed when surface, the area where the material is comthe material is being pulled off the spool and pacted onto the surface is heated. This is accomplished by blowing hot air at the nip point of the compaction roller (Fig. 22.1).
7 3 20 mm 0.13 mm
: 3.56mm 2.74 mm
I # of Samples. 1100
2.03
2.54 TowWdth (mm)
Fig. 22.4 Tow width histogram.
4.06
22.3 TOOLING CONSIDERATIONS
The tooling for fiber placement provides the shape of the part being fabricated and a suitable surface for the machine to lay tows onto. It also has features that can be used to associate the electronic data’s coordinate system, which describes the laying surface, to the machine’s coordinate system. The electronic data must accurately represent the finished tool. The smaller the surface of curvature is, the more important it is that the electronic data accurately represent the finished tool. If the electronic data does not accurately represent
480 Fiber placement the lay-down surface the tows will be placed in the wrong location and there is a chance that the machine could collide with the tool or part. It is also important that the electronic data is in a format that can be used by the offline programming system. If data is in the wrong format, the length of time it takes to program the part is increased and sometimes there is a loss of surface accuracy. If no electronic data exist then the tool needs to be accurately measured and this information needs to be converted into electronic data that can be used by an offline programming system. Tooling for fiber placement needs to meet criteria that are not necessarily required for other processes. The tool surface must have sufficient strength so that it does not deflect as the head compacts tows onto it. When the tool Fig. 22.5 Tooling features. surface deflects, the tows will not be placed in the correct position and they will buckle after the tool springs back into position. Tows that having to rotate the mandrel and they must are buckled will no longer carry the desired also be described in the electronic data coordesign loads. Another consideration is that the dinate system used by the part programmer. tool surface needs to extend 152 mm (6 in) past The probed coordinates are used by the offline the largest part boundary. This extended area software to create a transformation matrix is used by the head to roll off and onto the part that will convert the part’s electronic data as it is starting and ending a course. This coordinate system to the position of the tool in extended area also needs to be part of the elec- the machine’s coordinate system. tronic data, because the offline software also If the tool has to be removed from the uses this area as it is generating the courses machine and then replaced in after each part, when deciding where the ply boundaries are. it is important that the tool is designed with This extended area is also used as a place to features that allow it to be located in the same apply the bagging sealant tape for vacuum position each time it is put back in the debulking (Fig. 22.5). machine. This will eliminate repeating a time The tool must also be designed with fea- consuming alignment procedure. If the tool is tures that can be used to associate the part’s mounted between a headstock and tailstock it electronic data coordinate system to the is desirable to attach the ends of the tooling machine’s coordinate system. This can be shafts to the headstock and tailstock with nonaccomplished by perpetually molding or locking tapered adapters. This makes it easier machining three cross hairs onto the tool sur- to load the tool and when the tapers are tightface, in the area of the 152 mm (6 in) ened they are self locating. The tapered extensions. To provide accurate results, the adapters also need to incorporate a keyway on cross hairs need to be placed as far apart as the headstock end so the tool can be located in possible. These three cross hairs are probed by the radial direction. A tool mounted to a flat the machine to detect the location of the tool plate should be located by a three-point locatin the machine’s coordinate system, without ing system incorporating pins or keys.
Ply shape 481 22.4 PLY SHAPE
Fiber placement is used to fabricate simple parts such as flat panels or complicated parts such as an inlet duct, which is square on one end and round on the other (Fig. 22.6).The ply shapes can be any size or geometric form. A ply shape can also include interior ply boundaries which create holes. When generating ply shapes, the designer must consider the shortest tow length the machine can lay down. This length is the distance from the start of the lay-down point to where the tow is cut in the head. Figure 22.7 shows exterior and interior ply shapes of an
length. The exterior ply boundaries can the extended in the areas where there are minimum cut length problems. These extended areas can later be cut off. Interior plies must be reshaped to match the fiber angles. The designer must also consider the end locations of the tow. Each tow is cut at a 90" angle; because of this the ending angle of the tows may not match the angle of the ply boundary. Figure 22.8 shows three different ply boundary conditions for a 45" ply. The
Outer Pb bunday 45'
Inner Ply bunday -5.
Boundaries With Missing Tows
Adjusted Boundaries With No MissingTows I
I
I
Fig. 22.6 Fiber placing an inlet duct.
Fig. 22.7 Ply boundary adjustments.
aircraft structure. The black areas are the areas that cannot be fiber placed because the length of the required tow is less than the shortest allowable fiber length. These areas could be laid in by hand, or the ply shape could be adjusted, as in Fig. 22.7, so that the required fiber length " is longer " that the shortest cut
amount the tows can cover the ply boundaries is specified by the percent of ply boundary coverage. A 100% value is typically used on the larger ply boundaries that are later trimmed to shape. This makes sure that all of the fibers go to the edge of the boundarv. A 50% value is used on interior plv I
d
482 Fiber placement
PLY BOUNDARIES
0%
50%
100%
Fig. 22.8 Ply boundary conditions.
boundaries. The 50% value allows some of the fibers to go past the ply boundaries; but it makes sure that the correct volume of material has been placed in the ply. The percent of ply boundary coverage is specified in the offline programming software. 22.5 STEERING
The ability to steer the fiber band allows the designer to maintain a constant fiber angle on a complex surface, or align the fibers with some applied stress. Steering is made possible because of differential payout and because the impregnated tow has enough tack to overcome any sliding forces. Differential pay out of the individual tows is one of fiber placement's key features. This feature is required to allow each tow within the fiber band to maintain a unique length as the laminating action of the head pulls the tows off the spools and compacts them onto the part surface. When the head is required to laminate a curved path, the outer tows of the band will pull more length than the inner tows. Similarly, as
the head laminates over a contour, some tows must dispense more length than others. This allows the fiber placement machine to steer the fiber band. Steering can be defined as the binormal radius of curvature along a fiber path. An example (Fig. 22.9) shows O", 45" and 90" courses laid on a cone shaped mandrel. On the right side of the figure, the same courses are shown on the unwrapped mandrel surface. On the unwrapped mandrel surface it is easier to see which courses are being steered. The 0" course is straight showing that no steering is necessary. On the 45" course the amount of steering changes as the course is laid from the larger end to the smaller end of the cone. The steering radius on the larger end is 58.4cm (23.0in). It decreases to 20.3 cm (8.0 in) on the smaller end. On the 90" course the center line steering radius of the 2.54 cm (1.0 in) wide course is 28.7 cm (11.3 in). The 90" course also illustrates the need for differential payout. The fibers on the inside steering radius are 32.14 cm (12.653 in) long. The fibers on the outside steering radius are 35.11 cm (13.821 in) long. This is a 2.97 cm (1.168 in) difference in length. Tape is typically laid along it's 'natural path', but can be steered by small amounts. A tape laying machine can steer a 7.6 cm (3 in) wide tape along a 20.3 m (800 in) radius with small amounts of buckling. The buckling occurs because the fibers on the outside steering radius are in tension and the fibers on the inside steering radius are in compression. A typical fiber placement machine using 3.2 mm (0.125 in) wide materials can steer a fiber band along a 63.5cm (25in) radius without buckling the individual tows. When steering a radius smaller than 63.5 cm (25 in), the tows will begin to buckle if laid on a flat or a convex surface or 'Venetian blind' if laid on a concave surface. 'Venetian blinding' occurs when the fibers on the inside steering radius of the individual tows are adhered to the surface and the outside steering radius fibers are in the air.
Dropping and adding tows 483
'
0 DEGREE COURSE ON MANDREL
0 DEGREE COURSE
FLAT LAYOUT
p., /STEERING:
A5
58.4 R
DEGREE COURSE ON MANDREL 45 DEGREE COURSE FLAT LAYOUT
STEERING: 30.0 R LENGTH: 35.11 cm STEERING: 27.4 R LENGTH: 32.14 cm
ON MANDREL
90 DEGREE COURSE
FLAT LAYOUT
Fig. 22.9 Steering.
22.6 DROPPING AND ADDING TOWS
To start a course, the delivery head is positioned to the start point and the tows are fed out under the compactor. While following the fiber path, the delivery head can change the fiber band width by adding or dropping individual tows. This feature allows all or any combination of tows to be removed or added to the band width in increments equal to one tow width. With this capability, it is possible to decrease the band width to prevent adjacent courses from overlapping each other. A course is ended by cutting all of the tows as
the compactor lays the ends onto the surface of the mandrel. Each tow is cut at a 90" angle. When adjacent courses are placed next to each other, this can leave a triangular gap or overlap (Fig. 22.10). The designer can use the off-line software to specify the percent of overlap between courses. A zero percent specification will leave a triangular gap between courses. In this case the outside tow of the course being laid is dropped when it touches the adjacent course. On the other extreme, a 100% overlap will leave a triangular overlap and no gaps. The 100% overlap specification can leave a part
484 Fiber placement
TRIANGULAR GAP
\
/'
,' /'
,'
'.
'. '.
TRIANGULAR OVERLAP-/'\
0% OVERLAP
a u 50% OVERLAP
/
u
100% OVERLAP
Fig. 22.10 Overlap between courses.
'bumpy' because of the overlaps. The 50% specification is most commonly used because the smaller gaps and overlaps tend to average out each other as the laminate is built up. The other technique that helps average these triangular gaps and overlaps is to offset each ply by one and a half tow widths. This ensures that the small triangular gaps and overlaps do not align on top of each other.
which is to be manufactured by fiber placement. The first is concave surfaces and the second is areas with small radii of curvature. When consideringa part with a concave surface area, the designer must make sure the fiber placement head can fit into the concave area without hitting the surface of the part. There are some techniques that can be used to overcome some of these limitations. To help the head fit into a concave area the offline software has a feature known as collision avoidance. 22.7 SURFACE GEOMETRY The software knows the part and head geomeFiber placement can automate the fabrication try. It constantly checks to see if the two are of many composite part geometries that in the colliding. If they collide, the software will rock past could only be laminated using hand lay- the head off the surface normal away from the up. The types of surface geometries that can be collision. Figure 22.12 shows a part with a fiber placed range from fan blades to full 360" small concave area being fiber placed. As the asymmetrical shells (Fig. 22.11). There are two head stays normal to the surface and surface features that the designer needs to pay approaches the concave area, it will collide into special attention to when designing a part the part, as shown in the upper exploded view.
Surface geomety 485
F Fig. 22.11 Fiber placing a fan blade.
2
SURFACE NORMAL
COLLISION POINT
NO COLLISION AVOIDANCE
SURFACE NORMAL
7
C"
'-,a I-
I O " COLLISION AVOIDANCE
Fig. 22.12 Collision avoidance.
1
486 Fiber placement
To correct this problem, the software rocks back the head 10" off normal, as shown in the lower exploded view. There are limits to how much the head can be rocked off the surface normal. If the head hits on both front and the back sides, the software cannot avoid the collision and the area should be redesigned. Rocking the head to the front and back slightly affects the effective applied compaction force and the minimum cut length. Rocking the head sideways also affects the effective applied compaction force and requires extra compactor compliance. Examples of small radii of curvature are shown in Fig. 22.13. On the convex surface the compactor cannot completely come in
TOOL SURFACE
22.8 INSPECTION
Fiber placement is a very repeatable process and requires only a small amount of in-process inspection, but it is important to perform a rigorous first article inspection. This inspection is used to verify that the part program fabricates a part that meets all of the design requirements. This detailed first article inspection should not need to be repeated unless the part program has been changed. The first article inspection should start with a dry run of each ply. Dry running a program means to run the machine through the program without laying down tows. Dry running of a program verifies that the electronic data
7
SMALL CONCAVE RADIUS
SMALL CONVEX RADIUS
Fig. 22.13 Small radii of curvature.
contact with the surface. The part can still be fiber placed but the number of tows in the fiber band should be reduced to match the compaction line width. On the concave surface the radius should be increased to allow the compactor to contact the surface, otherwise the tows will bridge over the small radius of curvature.
describing the tool surface matches the actual tool surface. It also verifies that the transformation matrix, used to associate the part's electronic data coordinate system to the tool's position in the machine's coordinate system, is correct. This is accomplished by watching the compactor as it follows the part surfaces. Dry running also verifies that collision avoidance worked properly.
Further reading
487
Table 22.1 Inspection criteria
Discrepancy
Criteria
Correction procediire
A gap that is greater than 2.54 mm (0.1 in) wide and is longer than 76.2 mm (3.0 in).
Add a tow to fill the gap.
Wandering tow at start and end of course
A wandering tow that leaves a gap greater than 2.54 mm (0.1 in) wide and is longer than 50.8 mm (2.0 in). If the tow is captured under another tow, do not repair the captured tow.
Pick up the tow and reposition.
Twisted or folded tow
A twisted or folded tow that is longer than 76.2 mm (3.0 in).
Remove twisted or folded tow and replace.
Wrinkles at start of course A wrinkled tow that is wrinkled more than 19.05 mm (0.75 in).
Pick up the wrinkled end of the tow and straighten it out.
Bridging of tows
Use a heat gun and hand pressure to remove bridging or vacuum bag the part for a maximum of 15 min at full vacuum.
Bridging is more than 1.57 mm (0.062 in) high.
The next step in first article inspection is to load the machine with tows and fiber place each ply. After each ply is fiber placed, the ply is inspected for the correct fiber angle, ply location, band-to-band overlap and for missing tows. The easiest way to inspect for fiber angle and ply location is to have a Mylar template that has the ply boundary plotted onto it and a line for the fiber angle. The template needs to be located with alignment marks or pins that are part of the tool. A Mylar template will not work on surface geometries with curvatures in both directions. For these geometries, a formed inspection tool should be fabricated. After the first article inspection, subsequent parts need to have each ply visually inspected for excessive gaps and overlaps, lost tows, twisted tows, wrinkled tows and bridging tows. Table 22.1 is an example of a typical inprocess inspection criteria. Another criterion used to evaluate gaps is to take a 30.5 cm (12in)
scale and lay it on the ply normal to the fiber direction.All the gaps along the 30.5 cm (12 in) distance are measured and summed. If this value exceeds a pre-determined design criterion the area is reworked. FURTHER READING Barth, James R. 1990. Fabrication of Complex Composite Structures Using Advance Fiber Placement Technology. 35th Intern. SAMPE Symp., 2-5 April 1990. Enders, Mark L. and Hopkins, Paul C. 1991. Developments in the Fiber Placement Process. 36th Intern. S A M P E Symp., April 1991. Enders, Mark L. 1991. The Fiber Placement Process. Intern. Conf: Comp. Mater., (ICCM/8), July 1991. Evans, Don O., Vaniglia, Milo M. and Hopkins, Paul C. 1989. Fiber Placement Process Study. 34th Intern. S A M P E Symp., 8-11 May 1989. Evans, Don 0. 1993. Design Considerations for Fiber Placement. 38th Intevn. S A M P E Syrnp., 10-13 May 1993.
PULTRUSION
23
Brian A. Wilson
ment and manufactured the equipment to produce structural elements by the method. A The word pultrusion is used to describe a comtypical pultrusion machine is shown in Fig. mercial fabrication process for the production 23.1. The process has a relationship to extruof fiber reinforced composite elements. First sion, which is used primarily with metals and mention of the process is recorded in a patent describes the process of forming a shape using in 1951 with much of the early work in the a closed die and pushing normally hot metallic 1950s attributed to W. Brandt Goldsworthyl. materials through the die. Pultrusion differs in He performed much of the process develop23.1 INTRODUCTION
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Fig. 23-1Typical Pulkusion machine. (Courtesy of W. Brant Goldsworthy and Associates Inc.)
The basic process that it takes advantage of the strength of the longitudinal fibers in the section to pull the extruded shape through the forming die and the heated curing die to create a constant cross section structural element from a composite materials system. Hence the name pultrusion. Use of this-processhas now expanded both in the USA and around the world with many manufacturers equipped to produce the simple structural elements which are the main pultrusion products. The process was labeled for the first 30 years or so of its existence as a commercial production method. This was owing to the nature of the method, using longitudinal fibers and a simple thermoset resin system to produce a structural shape which had its primary strength in the longitudinal direction and properties in the transverse and third axis relating only to those of the resin system. The ability to incorporate three-dimensional strength first occurred in the mid 1960s when it became possible to include layers of mat/fabric and circumferentially wound layers of fiber both w i t h the body and on the surface of the structural shapes. Multi-directional reinforcement was used extensively starting in the 1970s2,3 . These developments, plus the introduction of thermoplastic resin matrices, have brought the pultrusion process into the manufacturing arena of the aerospace, defense and aircraft businesses as a relatively inexpensive and repeatable method of producing a constant cross-section structural element48 *. In this chapter, the basic process, variations on the process, pultrusion equipment, materials including fiber, fabric, resin matrices, additives, tooling, curing, structural connection methods, equipment manufacturers, pultrusion fabricators and potential markets and applications will be reviewed"". 23.2 THE BASIC PROCESS
The basic pultrusion process as originally conceived consisted of creating a cured composite shape by pulling a bundle of resin impregnated fibers through a heated die, providing sufficient
489
curing energy in the heated die to cure the composite into a hard structural product and using the resulting shape as a unidirectional strength element for commercial structures. The primary advantages of pultrusion are as follows: 0 0 0
~roductioniscontinuous~ labor requirements are low; material scrap rate is low; the requirement for support materials is eliminated, i.e. breathers, bleeder cloth, separator film, bagging film, edge tape.
23.2.1 VARIATIONS IN PROCESSING
The original process of pulling a bundle of unidirectional fibers through a curing die remained without variation until the middle of the 1960s. At this point, pultruders, both in Europe and the USA, developed the process of adding fabric in strip or sheet form and fiber mat to the pultrusion system in order to provide transverse strength and shear strength in the corner sections of complex pultruded shapes. In addition, the concept was developed for a filament winding head to be added to the pultrusion machine to provide a hoop wrap around a pultruded form. This process is known as 'pull winding' and is primarily used with tubular or rod shapes. It is not feasible to add the filament winding process to a structural shape that has a concave curvature, such as angles, channels and I-beams. A typical process of pull winding is shown schematically in Fig. 23.212. During the 1970s, the concept of pull forming was developed by Brandt Goldsworthy. Three versions of this forming process were developed. The first was actually named 'curved pultrusion' by Goldsworthy Engineering. The method was developed under a NASA contract and resulted in a pure, constant radius section with a constant cross section. A curved die was used, with a reduced radius, smaller than the required part radius. This allows for a degree of
490 Pultvusion COUNTER-ROTATING MANDREL
AND
\
LAYER 1. FIBERS
LAYER 2. WOUND FIBERS
LAYER 3. FIBERS
LAYER 4. LAYE&S. FIBERS WOUND FIBERS
Fig. 23.2 5-Layer tube on double head pull winder. (Courtesyof Pultrex Ltd.)
spring back after the product is removed from response to a specific production requirethe diel3. The die was split along its length, in rnentl4,l5.The pull forming process still uses the vertical plane. One half of the curved unit the basic principles of pultrusion since it has a was fixed in place. The internal radius portion continuous strand of fiberglass roving within of the die was further split in half vertically, the product and this roving is used to pull the along the central radius plane, forming two product through the sequence of the forming quadrants. Each of the quadrants had a slightly process. Goldsworthy Associates have longer circumferential length than the fixed designed two versions of this process, a half. The two moving quadrants moved past curved pull forming and a straight pull formthe fixed part of the die at the processing speed ing. These two versions of the process are of the pultrusion. When each moving quadrant briefly described in the following paragraphs had exited the plane of the fixed die, it moved but it should be noted that a major difference rapidly in a circular fashion to contact the end between pull forming and pultrusion is that of the following quadrant as it too passed pultrusion is a generic process which can be across the face of the fixed die. used to make many different cross-sectional shapes and products by changing the die in the machine (which creates the cross-sectional 23.3 PULL FORMING shape) and the materials of fabrication. Pull The pull forming process is a highly sophisti- forming however, is essentially a custom cated variation of the pultrusion process in process which must be tailored to a particular which both curved and straight product sec- product design. tions are fabricated with the ability to change cross-sectional dimensions of the product. It can be thought of as a combination of pultru- 23.3.1 CURVED PULL FORMING sion and compression molding. The pull In the curved version of the pull forming forming process was developed by process, a curved structure is manufactured Goldsworthy Associates and was created in using a selected combination of roving and
Pull forming 491 mat/fabric to satisfy the performance requirements of the produ~t'~J~. The die is a multicavity die. Several of these dies are mounted on the face of a wheel and are curved with the radius of curvature of the die matching the radius of its position on the wheel. The die is open faced and is also open on the two ends for entry of the fiber preform. The preform is impregnated with resin in the same manner as for pultrusion and the fiber and mat/fabric combination is pulled into the curved mold or die by the rotation of the wheel. As the preform is pulled into the die, a stainless steel band moves against the open face of the die and compresses the impregnated preform into the contour which is in essence a compression molding cavity. After the die is closed, it is heated to cure the resin and following cure, the stainless steel band peels away from the face of the die/mold. A fixed pin which is mounted adjacent to the wheel strips the product out of the mold and into a storage bin. When the stainless steel band is moved against the mold it is held in place by an automated clamping system until the product is fully cured at which point the clamping system releases the band which then retracts into a 'parking' position. The
machine is shown in Fig. 23.3. It should be noted that the process does differ from pultrusion in that the fiber preform is placed into the mold rather than being pulled through it. Also, while the cross section of the mold can change, the resulting part is a constant volume design, equivalent to the volume of the preform which is initially compressed into the mold. 23.3.2 STRAIGHT PULL FORMING
Straight pull forming was also invented by Goldsworthy Associates in response to a specific customer requirement for the automated manufacture of hammer handles. In this process, the fiber is pulled as a preform through the automated machine. Many single cavity dies are mounted on a belt. As the fiberglass roving moves through the system, a section of bulk molding compound (BMC) is cut from a rope of this material, is opened up longitudinally and placed around the roving and clamped onto it. The BMC is then compressed in a small briquetting press which travels with the fiber preform as it moves through the system. Following the compression
Fig. 23.3 Curved pull forming process. (Courtesy of W. Brandt Goldsworthy and Associates Inc.)
492 Pultrusion molding of the BMC, the press releases and the process variation has not been successful to balance of the fiberglass roving is fed in and date and further development will be required. encapsulates the BMC material. Following the encapsulation, a shrink film is placed around 23.4 PROCESS EQUIPMENT the wet package and the process goes through a heated tunnel which shrinks the film around The pultrusion fabrication machine consists of the product. Finally, the product is run through six different parts (Fig. 23.4): the creel; the the frame of a C-shaped die in a press which resin bath; the forming die; the heated curing provides the final compression molding and die; the pullers; the cut-off saw. curing of the hammer handle. The product is pulled from the die and cut off and the dies con23.4.1 CREEL tinuously exit the belt and are recycled back to the front end of the process. The creel is the beginning of the pultrusion This type of pull forming is a highly auto- process and is the material storage system from mated, sophisticated process which requires which the fibers and mat, veil or fabric are considerable capital investment for the design, drawn in the correct sequence to match the manufacture and set up. Each potential prod- design requirements of the structural shape. uct, which would be a candidate for the pull Since pultrusion is a long run continuous forming process, must have its design analyzed process, fiber rovings are provided in the maxand the process modified to respond to the cus- imum size configuration possible. Continuous tom requirements of the particular product. glass rovings are normally provided in 'center With the increasing popularity and explo- pull' packages between 14 and 23 kg (30 and 50 ration of the resin transfer molding process in lb) in size. These center pull packages are typirecent years, the pultrusion process has been cally stored on a bookshelf style creel. These reviewed for the potential of developing a creels have from three to six shelves and are repeatable, precise cross-section of a preform capable of storing anywhere from 45 to 120 using a loose fiber bundle and an adhesive packages of this type of fiberglass. These creels binder in place of the resin system. This are normally mounted on casters to create a Mat Roving racks
racks A
Material
Pulling mechanism
V
Preforming guides
Resin tank
Cutoff saw
Finished product
Fig. 23.4 Schematic of pultrusion process. (Courtesy of W. Brandt Goldsworthy and Associates Inc.)
Process equipment 493 mobile system. The casters are provided with a foot locking device to enable them to be locked in place when this is required. The glass roving is pulled vertically from the package through ceramic-lined holes in the shelves above. The roving is collected above the creel and turned 90" by means of a ceramic textile type thread guide and then moved forward to the material accumulating section just prior to the resin bath and forming die. A series of ceramic guides or rollers can be provided in the fiber path in order to tailor the tension in the fibers to the required level. The pulling of fiber from the center of the package will automatically insert a twist in the fiber as it is led into the pultrusion machine. To eliminate this, some fiberglass rovings are provided in a center pull twistless condition where the natural twist has been off-set by a 'built-in' reverse twist. Continuous fibers of fiberglass, carbon, aramid and polyethylene are normally sup-
plied on 'way wound' packages on standard diameter cardboard cores. A typical carbon fiber package mounted in a payoff creel is shown in Fig. 23.5. These fiber packages are designed to provide fiber payoff from the exterior of the fiber package and hence avoid twist. While fiberglass is normally supplied in the heavy center pull spools, it can also be obtained in the outside payoff type package. All of the cardboard cores are a standard three inch diameter with the exception of the aramid which is 90 mm (3.5 in) inside diameter. This style of package requires the use of a multiple spindle creel design in which the packages are oriented normally horizontal but usually with a slight upward camber. Creels are available with package positions up to 1500 or 2000 on a single creel or combination of creel frames. A loaded multiple spindle creel is shown in Fig. 23.6. Ceramic guides are used to move the fibers to the front of the creel
Fig. 23.5 Carbon fibre spool on package holder with quick braking action. (Courtesy of Texkimp Ltd.)
494 Pulfrusion (12-24 in) on cores of 75-100 mm (3-4 in) in inside diameter. This special mat/fabric creel must be able to accommodate rolls of these dimensions and resulting weight and permit the rolls of material to be installed in a varying sequence of locations to match the design of the final structure. This type of material creel will normally provide horizontal feed. However, if vertical mat or fabric is required, then an independent custom unit must be provided. These units are usually of a carousel-type configuration. Figure 23.7 shows a typical mounting for feeding veil rolls into the system.
Fig. 23.6 Loaded multiple spindle creel. (Courtesy of Texkimp Ltd.)
and into the pultrusion system. Tension requirements in this type of system are usually provided by spring attached tension straps which rub on a pulley attached to the package holder. This provides a braking action. The straps can be either independently loaded with weight or can be connected to a central I mechanical system, such that tension for all of the spools can be varied by a single adjust- Fig. 23.7 Feeding external veil material into pultrument handle. Fiber packages for this type of sion system. (Courtesyof Creative Pultrusions Inc.) system are normally 5 kg (11 lb). With this small size of package, it is normal for a large Clearly, the overall creel system for materials number of packages to be located on this type supply must be able to provide any combinaof creel assembly. In addition, package simula- tion or arrangement. It should be noted that in tion bars are normally provided with these every fiber and fabric system used for pultrutypes of creels to give a uniform tension and to sion, there should be sufficient continuous eliminate the possibility of a fiber from a small roving in the system to sustain the required or 'almost used' package touching and abrad- pulling force. As the various materials travel ing a near-by forward mounted, full package. forward towards the resin application area it is Stationed immediately after the fiber creel is important to control the alignment of the varithe specialized type of creel which is designed ous fibers and fabric/mat strips which are to accommodate rolls of mat, veil or fabric. going into the configuration. This will prevent These materials are usually supplied in roll knotting and twisting of the fiber reinforceform with diameters between 305-610 mm ment and also will ensure that the various fiber
Process equipment 495 materials remain in the correct relationship to each other and are placed in the correct zone of the pultruded product according to the product design requirements. This can be best accomplished by the use of orifice plates, creel cards combs and rollers (grooved and flat) to precisely and accurately place all the materials. The materials commonly used for these guides and rollers are primarily titanium oxide ceramic, stainless steel, tool steel alloys and chrome plated steel. These materials are also used for the various rollers associated not only with guiding the sheet materials but also with controlling them during their passage through the resin impregnation zone. 23.4.2 RESIN IMPREGNATION
Virtually all pultrusion processes utilize a resin impregnation bath to facilitate the impregnation of the resin into the fiber structure. The position of the resin impregnation bath in the production line can be varied and the manner in which the resin is applied to the fiber can have many different versions. A resin dip bath is most commonly used. During this process the fibers are passed over and under a series of rollers or bars which both spread the fiber to more easily accept the resin and also provide a massaging effect which 'kneads' the resin into the fiber bundles and structure. The wetting speed of the fibers depends upon their pretreatment and on the resin formulation. Wetting is also affected by the type of sizing agent on the fiber, the possible presence of remaining lubricant on woven fabric and finally the type of binder which is used in mat and veil products. The resin bath is uniformly used for products that utilize all roving in their construction or for products that are easily formed from the flat fiber ply which emerges from the resin bath. However, in many of the more sophisticated products which are now made from pultrusion, it is impractical to dip all of the materials in the resin bath. When vertical mats are required or hollow profiles are produced, a tailored resin
bath is frequently used which matches the preform shape or orientation of materials which pass through it. These types of tailored channels or baths can also be used to orient the flat materials properly. This method permits the resin impregnation to take place without moving the reinforcements away from their optimum path or shape. 23.4.3 VERTICAL PULTRUSION
The vertical pultrusion process should be mentioned at this point since the primary difficulty in creating a vertical pultrusion is the placement and use of a resin bath. In the arrangement for vertical pultrusion the equipment used is essentially similar to that for the horizontal methodI8.Most equipment operates in the vertical position with the exception of the resin bath and roving creels which are generally positioned horizontally and the roving is fed in the regular manner, horizontally through the resin impregnation bath. This bath is located above the entrance to the forming die and the fibers exiting from the bath are turned 90" across a roller and then proceed vertically into the forming die. The advantage of using this vertical procedure is that a uniform arrangement of impregnated fibers can be achieved across the section being formed. The effects of gravity are removed from the fiber arrangement. During the horizontal process, gravity plus any nonuniform tensioning across the fiber group will result in some fibers sagging under their weight with resin and hence not be properly located as they enter the forming and curing dies. An additional advantage of the vertical process is that it is possible accurately to locate the internal mandrels whch are necessary for hollow shapes and tubes. This is particularly important in the fabrication of heavy, thick wall tubes. In the horizontal process, these internal mandrels will frequently deflect under their own weight and cause a nonuniformity of wall thickness around the hollow profile or tube. In addition, the vertical process allows the use of several different fiber
496 Pul trusion entry points to the forming process with multiple resin baths. In the horizontal procedure, the multiple entry points have to come from the side of the main fiber path and it is not as easy to coalesce the fibers into a single merged preform. 23.4.4 USE OF THERMOPLASTIC MATRIX RESINS
Thermoplastics cannot be applied to roving easily using the standard resin bath, even with heating. Thermoplastic resins impregnated on roving are generally available as preimpregnated (prepreg) materials and are supplied by specialty companies who are experts in the prepreg process. Thermoplastic matrices improve the toughness of the composite and this is their major end use advantage. In general they have high softening temperatures and high physical properties coupled with a low fluid viscosity in their melted form. The prepregs are normally prepared using solvents and this provides some difficulty against environmental regulations in order to remove and dissipate the majority of the solvent. Some solvent has to be retained in the process in order to have the material be sufficiently pliable for machine and manual handling. In addition to the improved toughness of the thermoplastic composite, an important advantage of thermoplastic pultrusions is the ability to heat and reshape the product after f ~ r m i n g ' ~ - ~ ~ . 23.4.5 RESIN BATH LIFE
In a continuous pultrusion process, the pot life of the resin should be several days. However, if the volume of the resin bath can be kept small in relation to the resin being withdrawn, shorter pot-life resin, i.e. 3 4 h can be used. Shorter pot-life resins result in smaller batches and mixing becomes time consuming. 23.4.6 THE PULL WINDING TECHNIQUE
The pull winding process was developed in Europe and was used frequently by European
fabricators to provide pultruded tubular structures both round and square23,24. The process combines the standard continuous unidirectional fibers of the pultrusion process with hoop wound continuous fibers. The longitudinal fibers are used for axial and bending resistance while the hoop fibers are used for hoop tension and compression resistance. The combination of the two processes of pultrusion and winding provides virtually unlimited possibilities. However, the increasing complexity of the process limits these combinations. Normally, longitudinal fibers are positioned at the inside and outside surfaces of the tube with one or two hoop wound layers positioned internally in the tube wall. These hoop wound layers are not truly 90" hoop wound layers because of the slightly helical nature of the winding and the lateral movement of a wound roving, one band width of advance with each circumferential pass. Depending on the diameter of the tube, the winding angle is typically anywhere from 80-87". The winding can be performed in both clockwise and counterclockwise direction and in addition to the hoop tension and compression resistance can also provide a degree of torque strength to the tubing. Single and double head pull winders are used with these techniques. The pull winding process is shown in Fig. 23.2 and the procedure for the use of these systems is self explanatory. The important control feature of the pull winding process is the control of the winding speed and position with respect to the linear motion of the pultruded form through the machine. This is achieved by sensing the speed of the longitudinal motion accurately and using a microprocessor control to coordinate the speed of the rotating head motor. This relationship defines the pitch of the winding which is defined as the linear distance moved during one revolution of the head. With the microprocessor control the pitch can be varied, thus providing variations in fiber content and angular position, as required. A paper by D.E. Shaw Stuartz5
Process equipment 497 defines the primary advantages of pull winding over other methods of tube manufacture as: 0
0 0
it is a fully automatic continuous process; it is dimensionally accurate and repeatable; it produces tubes with a good external appearance and finish; it can have built-in color; it can be made with thinner wall sections than conventional pultrusion or filament winding.
used for forming dies include Teflon@,high molecular weight polyethylene, chrome plated steel and a variety of tool steel alloys. The fabrication of the forming die is a custom process which is best performed at the pultrusion fabricator by a skilled tool maker/machinist.
23.4.8 INJECTION PULTRUSION
In this modification of the resin impregnation process, the resin is directly injected into either the forming die or into the initial segment of the pultrusion die. The resin is injected into the 23.4.7 PULTRUSION DIES die under pressure and is forced into the interTwo types of dies are used in the pultrusion stices of the fiber system. The principal process: the forming die and the heating or curadvantage of this system is that it limits the ing die. Forming is normally accomplished release of volatile resin components and reacimmediately after the impregnation process tion products. although some shaping with the roving and An additional advantage of this process for mat/fabric inserts in a dry condition can take laboratory or research and development pulplace prior to entering the resin impregnation truders is that it enables a rapid resin change processing step. Forming dies or guides are without removal and cleaning of all of the normally attached to the heating or curing die resin bath components. In addition, dry fibers in order to provide the correct relationship are not impregnated with resin before entering between the forming and the heated curing the die and therefore they can be positioned step. For tubular or hollow profile pultruded accurately, even with complex profile cross products, a central mandrel support is necessections and multiple mat and fabric entries. sary internal to the fiber form and it is The schematic of the injection pultrusion sysnecessary to extend this mandrel as a cantilever tem is shown in Fig. 23.8. A typical resin through the pultrusion die. It is also important injection pultrusion die is shown in Fig. 23.9. to resist the forward drag on this central manIn addition to the advantages listed above, drel which will occur from the surface tension there are several disadvantages which must be and adhesive forces of the resin on the roving or weighed in a process trade off, before resorting mat. Materials must form in sequence around to the injection pultrusion method. These disthis mandrel and must alternate from one mateadvantages are: rial to another to prevent any weak areas at overlapping joints. The sizing of the slots, holes 0 a more complicated die design; and clearances in the forming die must be care- 0 more parameters to control; fully designed so that excess tension on the dry 0 analytical support is required to predict wet out behavior and design of the resulting or impregnated fiber is avoided. The fiber is dies; weaker in this condition than in the cured condition and filaments may be independently 0 fibers in the die are very tightly compressed and resin penetration is difficult, particubroken and distortion of the mat and fabric can larly with large wall thicknesses; take place. The forming die can also be designed to permit excess resin removal. This 0 the tightly packed fibers can act as a filter and partially remove resin additives, particprevents an abnormally high hydrostatic force ularly those in suspension. at the curing die entrance. Materials commonly
498 Pultrusion Fiber rack
doth racks
J
S -
I
-
\ / preforming
pulling mechanisms disengaged
\
Lc
I
hydraulic rams pressurized
guides
moving cutoff saw
t I
finished product
resin tank
Fig. 23.8 Injection pultrusion process. (Courtesyof American Composite Technology Inc.) injection port
thermostats
matrix injection fiber
+ I
‘1
i
6% zone
thermostats
1
cure zone
COmprBsEmn
zcne
Fig. 23.9 Resin injection pultrusion die. (Courtesy of American Composite Technology Inc.) 23.4.9 HEATING AND CURING DIES
There are three considerations in the design and use of a heated die for the curing of a pultruded composite section. The first of these is the positioning of the die relative to the machine access, the second is the actual method of heating to induce the necessary energy into the composite material to fully cure the resin system. The third includes the design features and construction materials for the die.
23.4.10 POSITIONING OF THE HEATED DIE
The heating/curing die must be firmly mounted to the frame of the pultrusion machine in order to react the pulling forces, which are created in pulling the product through the process. These forces are typically in the range of 5440-7250kg (6-8tons). However, some of the larger machines which are capable of producing parts up to 1.52 m (60 in) wide and heights of 0.60 m (24 in) (with narrower parts) can require pulling forces up to 16 320-18 000 kg (8-20 tons). The mounting method must not only permit these types of loading but must also be capable of mounting height adjustment in order to accurately align the axis of the die with the pullers which move the product out of the die. Following alignment and clamping, the mounting system must also not allow any deflection which might provide an angular or dimensional mismatch of the machine. 23.4.11 DIE HEATING
Die heating is probably the most critical control parameter of the whole pultrusion process. The die heating profile will control
Process equipment 499
the rate of polymerization of the resin system and the position of the resin gel point front w i t h the die. It also influences the degree of resin exotherm profile throughout the various wall thicknesses of the pultruded structure. The curing dies are typically heated with electric strip heaters or hot oil jackets. The thermal curing using these two methods of heating is slow, owing to the fact that the tool steel of the die is a poor conductor and limits the rate of heat injection to provide a uniform cure. The thermal energy which is required to cure the composite material must all be applied through the outer surface of the composite shape. The heat input is required to produce gelation of the resin throughout the composite part. When the composite structure being pultruded has thick sections, this necessarily requires a longer heat input which slows down the pultrusion rate. Increasing the rate of heating to permit a faster pulling speed will not necessarily solve this problem and may result in premature curing of the outside skin of the profile with early onset of the exotherm in the resin system and potential overheating and cracking of the exterior surface of the part. Temperature conditions of the die are controlled by the internal placement of thermocouples and this can result in individual control of heating rates in various segments of the die. Instead of using strip heaters or hot oil jackets, it is possible to heat the curing die by means of heated platens and a press. These heated platens will usually have several zones of heating control with imbedded thermocouples to sense the platen temperature. T h s method provides a uniform heating condition to the die. However, since the thermocouples are sensing the platen heating temperature rather than the die or product temperatures, the temperature differential between these zones must be well documented. This permits the platen heating temperature to be set to provide an adequate heating level in the composite product and absorb all of the conduction and radiation losses from the system.
Use of the platen system will allow change out of dies with very little system down time. During start up and also during shut down periods it is necessary to provide a cooling method at the front of the die to prevent early gelation of the resin system. During these phases, the entire die will heat up rapidly without the composite being pulled through it and acting as a heat sink. The cooling can be done dynamically through cold air impingement or by using a water jacket or tubes through which cold water will be pumped. Instead of these dynamic methods, a simple static method would be to provide an unheated zone at the beginning of the die to act as a heat sink and conduct away the excess heat during start up or shut down actions. These same unheated or cooled sections may also be useful at the exit of the die to remove contained heat from the product prior to its exit. It has been mentioned that the heating profile within the die is the primary factor which controls the speed of throughput of material in the process. This was recognized early in the development of the pultrusion process by W. Brandt Goldsworthy and he added a radio frequency preheated system to his pultrusion machine designZ6rz7. The use of radio frequency heating in conjunction with the traditional die heating methods can significantly increase running speeds. This technique is basically limited to unidirectional reinforced rods, bars and shapes and cannot be used if carbon fibers comprise the reinforcement. Data from Goldsworthy Engineering Inc. shows speed increases of up to 400% based on a 2.54mm (0.1 in) wall thickness pultrusion. Speed increases of 100% (i.e. doubling the speed) are accomplished at wall thicknesses of 2.54 mm (0.1 in). During recent years it has become possible to evaluate the internal die profile with regard to both temperature and pressurez8.The sensors which are used to provide that data are essentially similar to strain gages with output leads attached to them. The pressure sensor is
500 Pultrusion a unique development resulting from the heating zones in order to maintain control music industry. These sensors were originally over the resin gel, curing and cooling of the used as striking pads in electronic keyboards pultruded section within the die. In addition, and the manufacturer has now developed if one of the cavities in the multi-cavity tool is them to measure pressure forces transverse to damaged, the complete tool is not out of prothe plane of the sensor. The output from the duction since the remaining useable heating sensor is transmitted via a pair of thin strain zones can be modified to continue to provide gage type wires. It is thus possible to insert an acceptable product. The dies are usually made from steel, which both of these sensors (temperature and pressure) into the fiber resin system prior to can be hardened tool steel, or steel alloy which entering the curing die. The sensors will then is treated, hardened and plated before use. The travel through the die internal to the product die must be relatively thick walled so that it and will record both the temperature and the can be heated uniformly and retain the heat internal pressure in the die from the entrance input. Thick wall design also reduces distorto the exit. Once the sensors have traveled tion under temperature and pressure. It has beyond the exit from the die, the connecting been suggested that the cross-sectional area of wires can be cut. The section of product which the steel should be at least ten times that of the contains the two sensors and the lead wires is pultruded part. The steel used should be hard then cut out and discarded. The resulting data since it has to withstand the abrasive action of provides a complete process description from the composite being drawn through it. A typical hardness is a Rockwell C rating of 30, the entrance to the exit of the die. which can be obtained with a prehardened tool steel. Dies are manufactured from multi23.5 DESIGN FEATURES AND MATERIALS ple pieces that are machined and joined In considering the design of the heated curing together to create a design profile cavity. As die for a pultrusion system, it is possible to use the various sections of the die are connected a single cavity tool, a multi-cavity tool or sev- together they must be properly aligned. This eral single cavity tools mounted in parallel. can be done using dowels for permanent The choice between these three will depend alignment or by machining an alignment upon the size, complexity, dimensional toler- groove on the outside part of the die which ance and surface quality of the pultruded can then be used for a hard metal key which product. Individual or single cavity dies are a can be driven into the groove to provide a posfrequent choice of pultruders, particularly for itive alignment for the entire tool. The die medium to large or sophisticated shaped assembly is then match drilled and tapped for products. Use of the multi-cavity tool or sev- the assembly bolts which hold the parts of the eral single cavity tools mounted in parallel is die in position against the high internal presgenerally restricted to very competitive and sure developed during the process. Following low cost shapes such as round and square alignment and assembly, the exterior surface cross section rod. The choice between an indi- of the assembled die is ground flat. The intervidual die and a multi-cavity die is frequently nal cavity surfaces are polished using dependent upon the experience and prefer- polishing wheels and buffing compounds to a ence of the manufacturing engineer in the high surface finish 0.254-0.762 wm (10-20 kin). pultrusion company. The multi-cavity tool can At this point a bell mouth is machined be an efficient arrangement for production around the entrance of the die to provide a conditions. In this case the die is two to three smooth entrance for the resin impregnated times the length of a single cavity tool. The fiber form. The radius will vary in size from longer die makes it possible to use different a small radius for small simple structural
Design features and materials 501 profiles to a relatively large radius for a large and complex composite profile particularly where the fiber content is high. The bell mouth is polished and the finished die is hard chrome plated to a thickness of 0.025-0.050 mm (0.0014.002 in) with a Rockwell C hardness of approximately 70. The dimensions and surface quality of the pultruded product are a direct reflection of the condition of the die. Dies will thus not be removed from production to be inspected unless the shape or quality of the product begins to change. Chrome plated dies will normally have a life of 61 000-150 000 m (200 000-500 000 ft) of product run with their initial chrome plate application. Up to 305 000 m (1x lo6ft) of product have been produced from some hard chrome plated dies. Chrome plated dies require frequent inspection to insure that their internal shape and dimension is maintained. They should also be inspected for wear of the chrome plated surface since the wear process will proceed much more rapidly if the tool steel surface is exposed from under the chrome. Once the die has worn and produces product beyond allowable dimensional tolerances, the die may be replated and repolished back to original dimensions. This process may be repeated several times.
pultruded profile remains stationery until the puller returned to the beginning of its stroke. Because of the alternating pull and pause mechanism this system was known as the 'intermittent puller'. This system is still used on some very early machines, however, it is certainly not in widespread use. A modificationof the clamping/pulling system has become popular which provides a continuous pull. The clamping, pulling and unlocking cycles of this system are coordinated by the control system of the machine. The drive system used can be either a hydraulic cylinder, a threaded ball screw such as is used on lathes, or a chain drive mechanism. The clamping pads are still formed to match the contour of the profile being pulled. The two puller heads must operate in the space originally designed for the single puller. Consequently, limited lateral movements of approximately two feet each are sustained by the two pullers. The two puller system is shown in Fig. 23.10.
23.5.1 CLAMPING AND PULLING
Three different types of clamping and pulling systems have been used in the pultrusion industry. Of these three, only two are now encountered. The original system used on the earliest machines in the 1950s and early 1960s employed a single clamp. This clamp was hydraulically controlled and contoured pads were used for gripping the part. The unit containing this single clamp was pulled by a continuous chain for a distance of 3.2-3.7 m (10-12 ft). Areversible motor was used to drive the chain and following the pull stroke, the puller/clamp released the product and returned to the beginning of the pulling stroke, clamped the product and pulled againz9.The obvious disadvantage of this system is that the
Fig. 23.10 Conventional (Courtesy of Pultrex Ltd.)
two puller system.
Continuous belt pullers are used on basic commercial machines. These pullers can be used with single or multiple cavity molds. The cleated chain or caterpillar version of the continuous belt machine has many individually contoured puller pads attached along the chain length. The number of these contoured
502 Pultrusion puller pads depends on the complexity of the part but generally varies between 12 and 60 pads. This large number of pads permits a lower unit pressure between the clamping pad and the pultrusion. The caterpillar type machine was designed and built in 1958 by Brandt Goldsworthy, Dennis Franks and Tom Bailey. Caterpillar type machines are preferred and still widely used in the industry. 23.5.2 CUT-OFF STATION
Every pultrusion machine utilizes a cut-off saw to cut the pultruded profiles to the required length for shipment and use. This saw is frequently of a radial arm type as shown in Fig. 23.11, but can be also a chop saw, orbital or band saw. It is mounted on a platform which moves down the pultrusion exit table at the same speed as the pultruded product. Carbide and diamond tipped saw blades are used for the cutting of glass and carbon pultrusions. However, these saw blades are not effective for cutting aramid fiber pultrusions. This fiber is known for its properties of toughness and resistance to abuse. These properties lead to difficulty in machining, grinding and cutting. The most successful method of cutting
aramid fiber to date has been the high pressure water jet and presumablythis could be adapted for use on a pultrusion machine. The inherent disadvantage of the high pressure water jet is its cost which is from $50 000-100 000. A compromise solution to this problem would be to cut off the sections as smoothly as possible, using the diamond saw, to a slightly over length condition with very rough ends. The product length can then be subcontracted to a waterjet cutting source for final trimming. This will result in some wastage. 23.6 MATERIALS
Fiber properties to aid the designer are shown elsewhere in this text. Following the selection of the fiber type to suit the required design factors, the fiber must be oriented in the correct direction. It is understood that all of the fiber types must be available in continuous form in order to be useable in the pultrusion process. The most commonly used form of continuous reinforcement is roving. This is available in single and multiple strand configurationsm.Glass rovings are designated by their yield which is the number of yards per pound of material. The two most commonly used versions are at 112yd/lb or 224 in/kg or 112 or 124 m/kg (56 1 or 62 yd/lb). The glass rovings are typically supplied in 18.1 kg (40 lb) hollow cylindrical packages with a center pull payout. A similar center pull spool is also available for both aramid and polyethylene fibers. Carbon fiber is typically available in either a 3K, a 6K or a 12K !!- filament. It should be noted that the tow sizes of 1 carbon are much smaller than the glass rov- the ing and package weights are 1-2 kg (2-5 lb) with an outside payoff designed for a package holder style creel system. New versions of the - carbon fiber roving are available now in 40K, 160K and 320K tows. Use of these tows allows the fiber to be laid down very rapidly and consequently these versions of carbon fiber are attractive to pultruders. Typical properties of Fig. 23.11 Conventional cut-off saw. (Courtesy of fibers used in PultruSiOn are shown in Table Creative Pultrusions Inc.) 23.13', 32.
\
Materials 503 Table 23.1 Typical properties of the major fibers used in the pultrusion process
Property
E-glass
S-glass
Keular (Aramid)
Density,
2600
2491
Tensile strength, MPa
3447.5
Tensile modulus, GPa Elongation at break, %
Spectra
Carbon (Type T300)
Carbon Inter. modulus
1470
968
1720
1770
4585.2
2964.9
1170.0
1896.1
2560.9
72.4
86.9
131.0
26.0
379.2
473.6
4.8
5.4
2.3
3.7
0.5
1.81
~~
kg/m3
All the rovings discussed will yield the highest possible longitudinal properties. Fibers as rovings result in the maximum fiber content to be achieved in pultrusion. If the longitudinal rovings are used under near perfect conditions, a 65% fiber volume percent level should be achieved. In a product which utilizes 100% roving this material is normally in the longitudinal direction or axis of the pultruded product. Properties in the other two directions are dependent upon the resin system and the mechanical properties of the matrix resins are much lower than fiber properties. Transverse strength problems are overcome by inserting transverse fiber materials into the pultrusion. This is done either by using fabric or continuous strand mat. The latter is most commonly used. While the fabric is a standard woven textile form, the continuous strand mat has fibers oriented in a random mode, bonded with a thermoset resin binder which holds the mat together adequately for processing in the pultruded section. While mat is available in any of the fibers which have been previously discussed, the most common available mat is an E-glass version that has coarse fibers in an open or porous construction. This mat can be used either as a center ply in a pultruded structure or on the outer surface of the structure. Use of the mat greatly improves the transverse physical and mechanical properties. It should be noted, however, that the mat
is porous and its use on the exterior surface of a pultrusion might well leave porosities or voids in the surface. To counter this problem, a very fine filament, E-glass mat, commonly known as veil, can be used as the surface ply. Its presence during the pultrusion process will tend to bring more resin to the surface of the pultrusion and this will achieve a smooth, uniform surface, devoid of porosity or voids. The veil mats can also be placed internally in the composite and recent improvements in their structural properties have made this possible. The random fiber mats in E-glass are used in weights of 0.15-0.6 kg/m2 (0.5-2 oz/ft2). The inclusion of these mats in the pultruded structure means that some of the longitudinal fibers will have to be removed to allow for the volume of the mat, veil or fabric. With the use of fabric or mats in the structure, the resin content by volume will increase in order to fill the openings in the mat or fabric. Thus while the transverse strengths increase, the longitudinal strengths usually decrease. Mats are also available in carbon fiber. The random fiber structure of strand mat provides fibers in all directions. However, this random orientation does provide some problems in that the fibers may not provide a symmetrical balance within the structure. The initial solution to this problem was the use of woven fabric. However, the lack of tension in the fabric results in a lower strength capability
504 Pultrusion of the pultrusion since under load, the fibers in the fabric will have to straighten and become tensioned prior to being able to accept load. One way of solving this difficulty is to use non-woven biaxial fabrics which are stitched or knitted together at the crossover points. Because of the nature of fabrication of these nonwoven materials, any ratio of fibers in the two directions can be provided. It is also feasible to utilize k45" fibers in conjunction with the 0 and 90" fibers. The biaxial fabrics are normally used as internal plies and not on the external surface. This is due to the nature of the nonwovens in that their transverse fibers will tend to be displaced by friction with the walls of the die during pulling. Fabrics using a +45" orientation without any longitudinal fibers are usually impractical for the pultrusion process. It should be noted that hybrid composites with tailored properties are possible using combinations of the fiberglass, carbon and aramid materials. The designer will readily determine the mechanical properties which are required from the nonwoven or woven fabrics. The rule of mixtures will apply for combination proper tie^^^. When pultruded composites are used in outdoor weather conditions, the surface of the composite may be degraded with time by sun, wind, rain and ultraviolet exposure. In order to solve this problem, additional resin needs to be provided at the exterior surface of the pultrusion. This is done by incorporating the very fine filament veil mats which are typically fabricated from polyester or nylon34. These veil fabrics are available in a variety of weights and weaving patterns. They help the pultruder by providing a tough surface material which will protect the die wall from the abrasive nature of the fiberglass or aramid. In addition, the resin rich surface is created without any obvious fiber weaving patterns, plus the veil materials can be screen printed with company identification or decorative effects35.
23.6.1 MATRIX RESINS
Of all of the technology considerations in the pultrusion process, the most critical material is the resin system and its f~rmulation~"~~. Resin selection controls mechanical characteristics, electrical insulation, corrosion resistance, operating temperature and flame and smoke properties. It also has a significant effect on the process speed because of the required cure cycle for any particular resin. The selection of a resin system will also affect the production cost of the process. The two most commonly used resin types in pultrusion are the isophthalic polyester and the vinylester. These two comprise over 90% of all resins used in pultrusion. Epoxy resins and phenolic resins are also being increasingly used. Phenolic resins were traditionally avoided by pultruders because of their condensation reaction during cure. Condensation reactions produce large volumes of water vapor and this typically causes voids, channels, delaminations and porosity when there is no provision to remove it. While the pultrusion die does have an entrance and an exit, nonetheless the system is essentially a closed, pressurized volume. Table 23.2 provides typical mechanical properties for resin systems most generally used in pultrusion and several other chapters discuss the broad range of matrix materials39.
23.6.2 PHENOLICS
In consideration of the use of phenolic resin systems, the disadvantage of the condensation type reaction was certainly sufficient to cause delays in the potential use of phenolic resin in the pultrusion process. The amount of water vapor which is generated in the condensation process is very large and it has always been assumed that a phenolic pultruded structure would look somewhat like a sponge. However, phenolic systems have been pultruded in recent years and a phenomenon has occurred which is not well understood. During pulling of a phenolic/fiberglass structure through the
Materials 505
Property
Polyester
Vinylester
Epoxy
Density, kg/m3
1100
1100
1300
Tensile strength, MPa
77.2
81.2
75.4
Tensile modulus, GPa
3.3
3.4
3.3
Elongation at break, YO
4.2
4.5
6.3
Flexural strength, MPa
122
134
115
Federal Department of Transport and the various state departments of transportation have imposed fire controls on composite materials. It is required that they will not bum or stimulate combustion, have minimum required smoke levels and also will not produce toxic fumes under flame impingement and high environmental temperature conditions. Considering all of the resins in the composite industry, phenolic resins will come closest to matching these flammability specifications. Because of these factors, it is anticipated that the use of phenolic resin systems in pultruded products in the future will increase dramatically.
Flexural modulus, GPa
3.2
3.1
3.3
23.6.3 COMPARISON OF RESIN SYSTEMS
Heat distortion, "C
77
99
166
Table 23.2 Typical mechanical properties for resins used in the pultrusion process
pultrusion die, a high pressure jet of steam is noted at the exit from the die. How and why the steam is caused to come off in this manner is not known. However, in the experiments which have been run, the resulting pultrusion has not had any porosity problems and the processing tests are noted as being successful. Specific pultrusion grades of phenolic resin systems are now available from plywood manufacturing corporations such as Weyerhauser and Georgia Pacific. The availability of these resins would certainly indicate that the anticipated processing problems have not occurred40, 41. The importance of phenolic resins is in their resistance to fire and their low smoke and toxicity production under fire conditions. All forms of composite materials, including pultrusions, are being used increasingly in mass transit, aircraft and civil engineering applications. In all of these areas of application, increasing contact of the composite material systems with the general public is occurring. Because of this, the fire smoke and toxicity requirements of specification control groups such as the FAA, the
In recent years, there has been use of epoxy and phenolic resins in pultrusion. G.A. Hunter of Shell Development Company compared the properties of resin systems42.He provided a three zone model of the pultrusion process within the curing die (Fig. 23.12). The sketch provides an excellent background for comparison of resin proper tie^^^. Of the four primary resin systems used in the pultrusion process, the polyester and vinylester resins account for more than 90% of the marketplace. Phenolic and epoxy resins make up the balance of the market. In comparing resin systems, one should review the internal contours and the heating profile of the heated resin die and examine the change in resin morphology as it proceeds through the die. The model of the pultrusion process given in Fig. 23.12 which shows the three zones of the heated die and the transition of the resin phase from liquid through the gel zone into the solid phase. The first zone shown in Fig. 23.12 is where the material enters the die at room temperature and expands as it absorbs heat which causes the hydraulic pressure in this zone of the die to rise. As the material progresses into zone 2, or the gel zone, it has absorbed more heat, is beginning to cross link and changes from a viscous liquid into a nonflowing jelly type of material, then into a
506 Pultrusion ~~
"THE GEL ZONE" STRIP HEATER
&&(
SOLID P H A S E
LIQUID PHASE
..
i.
ZONE 1 iVlSCOUS SHEAR FORCES
*........
i
i
ZONE 3 SLIDING FRICTION FORCES
............. 2
i
**a.
.*......-.I
:
.. . i
ZONE
I..
:
COHESIVE FORCES
Fig. 23.12 Three zone model of the pultrusion process. (Courtesy of Shell Development Company.)
rubber-like texture. As the material cures to a hard solid, shrinkage occurs which releases the hydraulic pressure forces and the product shape retracts from the internal surface of the die. This is zone 3. In this zone, because of the release of the product from the surface of the die, the sliding frictional forces are very slight. Depending upon the thickness of the part and the process speed, the bullet-nose shape of the
gel zone will expand or contract. Joseph Sumerak in 1985 quantitatively described the internal dynamics of the pultrusion process. Taking test results from Sumerak's earlier work, Hunter showed the relationship of pull loads to processing speed for catalyzed and uncatalyzed resin systems (Fig. 23.13)w7. For the uncatalyzed resin case, the rising pulling load associated with
looot7 1 P t W I T H 20 P H R C L A Y
D
800
*
700-
600
-
PULL LOADS ATTRIBUTED TO
VISCOUS SHEAR AND FRICTION O F
0
A I
-
-A
-
CATALVZED RESIN P U L L L O A D S A T T R I B U T E D TO PURE VISCOUS SHEAR OF THE U N C ATALYZ E D R E S I N CALCULATED PULL LOADS ATTRIBUTED TO VISCOUS
u)
S H E A R O F T H E CATALVZED
0 400J J
3 300-
n 200
-
0
12
LINE SPEED,
24
36
INIMIN.
Fig. 23.13 Pull loads compared with line speed for different types of resin systems. (Courtesy of Shell
Development Company.)
Materials 507 increased processing rate, or line speed, is the result of increasing shear forces over the length of the die. In the case of the catalyzed resin and referring back to Fig. 23.12, viscous shear forces are generated only in the front portion of the die, i.e. zone 1. Within the gel zone, cohesive forces come into play for a small length of the die which is followed by the transition to the rubbery cured material which provides substantial friction forces. As the resin hardens and shrinks away from the surface of the die, the frictional forces are reduced significantly. It is obvious that the pull load is significantly higher for the catalyzed resin system, particularly as the line speed increases. This proves that the major portion of the pultrusion loads are generated in the gel zone and are cohesive forces and frictional forces resulting from the interface of the resin and the die. Sumerak showed that a significant part of the internal pressure does develop inside the
B
---
-1-
4
0
>
INITIAL CURE CYCLE SECOND HEAT CYCLE
STARTING VOLUME
-2-
die and is proportional to the speed of processing. Hunter provided evidence that the pressure loss in zone 3 of the die occurs well before the material cools. Thus it is not thermal contraction but volumetric shrinkage due to the cure of the resin. The coefficient of thermal expansion of the steel material of the die also enters into this equation. For a differential temperature of 121°C (250"F),the hottest temperature section of the die for a 12.7 mm (0.5 in) diameter pultrusion will be 0.3% larger than the entrance. Thus pressure and volumetric shrinkage together play a major role in pultrusion dynamics. Insufficient pressure causes sloughing problems and insufficient shrinkage can cause excessive pull loads. The resin rate of shrinkage affects the rate of pressure decay and is controlled linearly by the cure rate of the resin. Thus a delicate balance between pressure, cure rate and shrinkage must be obtained for a clean pultrusion process to take place.
NET VOLUME LOSS
-
9.098 ML 0.<5
VOLUMETRIC SHRINKAGE -31
.
r
l
r
.
l
:
ML
1
6.04 % 1
1
.
1
I
,
Fig. 23.14 Volume change of polyester resin during cure. (Courtesy of Shell Development Company.)
508 Pultrusion 23.6.4 SHRINKAGE
Hunter ran shrinkage tests on a typical polyester resin and a standard Shell epoxy resin system42.The volume change of the polyester and the epoxy resin are during cure are shown in Figs. 23.14 and 23.15. The data shows that the polyester shrinks almost twice the degree of epoxy. However, Hunter reports that the net shrinkage is not nearly as important as the profile of that shrinkage. The polyester continues to expand after the gel point which is followed by a high shrink rate that gradually tapers off. In comparison, the epoxy resin shrinks before it gels and continues to shrink at a steady rate until it is fully cured. This information sheds new light on the understanding of the pultrusion characteristics of epoxy resins compared to polyesters. In most composite manufacturing processes, the application of pressure during the curing phase of the resin is always beneficial to the resulting product. Pultrusion is no exception to this rule and it is beneficial to generate pressure in the pultrusion die up to the point where the resin
is fully cured. Similarly, the pressure during the gelation phase ensures that the product is tightly held against the surface of the die and consequently a smooth surface will be generated with the pressure preventing sloughing. Thus, from a comparison of the test results, it is obvious that the polyester shrinkage profile is superior to the epoxy in terms of providing gelation and cure under pressure. In addition, the sudden high initial rate of shrinkage following gelation for the polyester resin is also beneficial in that it results in a fast pressure drop and hence frictional force reduction. In comparison, the epoxy resin begins to shrink well in advance of gelation and gels under a condition of declining hydraulic pressure. Thus much of the hydraulic pressure is lost by the time that gelation occurs due to the effect of volumetric shrinkage. Following gelation, the rate of shrinkage is very slow such that it causes the friction forces to only reduce gradually. Thus in the gel zone for epoxy resins there is insufficient hydraulic pressure to prevent sloughing. This explains why a
3t. A GEL POINT
i
22.8" C QUENCH
-2-
NET VOLUME LOSS
-
.32 MLS.
VOLUMETRIC SHRINKAGE
-3-
-
3.57 %
Fig. 23.15 Volume change of epoxy resin during cure. (Courtesy of Shell Development Company.)
Materials 509 problem is frequently encountered when epoxy resin is substituted directly for a polyester without consideration of the curing and shrinkage properties. Hunter postulates that a simple solution has been found to compensate for these shrinkage characteristics in epoxy resin systems. The presence of fillers, whether fiber or powdered, in the resin system, reduces the amount of total volumetric shrinkage. Also the pressure from thermal expansion is directly proportional to the amount of filler or fiber reinforcement volume. Thus the increase in reinforcement to resin volume ratio, either by fiber or powder fillers, will reduce the shrinkage tendency and the hydraulic pressure will be increased at the same time. Even though the epoxy resin shrinks prior to gelation, sufficient pressure will remain to prevent sloughing. This explains why it is always beneficial to have a higher fiber to resin ratio for epoxies in the pultrusion process than for polyester resins.
23.6.5 CURE RATE
As previously mentioned, shrinkage rate is a direct result of the known cure rate of the resin. It is beneficial in the pultrusion process to have a high shrinkage rate to initiate a quick pressure drop to reduce frictional pull loads. From this point of view, it is important for the epoxy to have a fast cure rate. This will also provide a shorter gel zone which will result in a faster processing rate. Cure rates of polyesters may be varied chemically by changing the amounts and types of peroxide catalysts which are used to initiate them. It is not simple to change the cure rate of an epoxy resin chemically. Curing agents for epoxy resins are selected based on the desired performance parameters for the epoxy in the final product or structure. Some considerations of pot life and manufacturing process also influence this selection. Thermal accelerators can be used. However, the effect of increasing cure rate accelerator
00
50
00
50
75
I
I
I
I
100
125
150
175
200
GEL TIME TEMPERATURE, DEG. C
Fig. 23.16 Gel time compared with temperature of epoxy and polyester resin. (Courtesy of Shell Development Company.)
510 Pultrusion 23.6.6 REINFORCEMENT VOLUME
content may be to reduce the pot life. Hunter ran experiments to provide data on gel times of epoxy resin at different accelerator levels. In addition, he checked on the gel times of polyester resin using two different curing agents. Figure 23.16 shows the data which resulted from these two evaluations. The data on the epoxy resin system shows that significantly more heat is required to generate the same gel times even though the accelerator content is doubled. Figure 23.17 shows the viscosity versus time of the epoxy resin at two accelerator levels and two temperatures. This graph clearly indicates that pot life is sacrificed by increasing the accelerator level. Pot life is also affected by temperature and Fig. 23.17 illustrates that a small increase in temperature will reduce the time to double the initial viscosity by almost half. Heat can be generated during the mixing process of the epoxy formulations and because of this it is important to minimize the mixing times when using high shear mixers that generate heat within the body of the resin system. The implication of the experimental data presented in these figures is that the most efficient method of increasing cure rate in the epoxy resin is to increase the die temperature. 4 0 0 0 ~0- P T D . 3500
A C C E L E R A T O R
The general relationship of glass fiber content to pull loads in epoxy resin system is shown in Fig. 23.18. These data were derived from an experiment in the Shell laboratories where the reinforcement volume was decreased and the pull loads were recorded until sloughing The fiber volume was then increased until the sloughing was eliminated and was increased further until pull loads became too large. The data shows that there is a plateau in the pull load curve spanning approximately 2% of the glass fiber content range. This is the optimum level for pultruding the 12.7mm (0.5in) diameter epoxy rod used in the test program. Below the optimum range, sloughing occurs owing to the insufficient hydraulic pressure at the gel point. Above the optimum level, the pull loads rise owing to the high pressure both during and after the gel zone (referring back to Fig. 23.12). Both polyester and epoxy systems respond similarly to the different types of reinforcement materials which are contained within the pultruded structure. For both of these resin systems, the minimum reinforcement level to prevent sloughing when using a continuous mat and roving is somewhat less than that for an all rov-
LEVEL AT
2 l . C
-
W u)
0
3000-
n 25002
w
0
>.
2000-
1500-
u)
2
-
1000-
u)
5
50001
I 0
I 1
I 2
I 3
I
I
1
1
1
1
4
5
6
7
8
9
1 1 I J 1 0 1 1 1 2 1 3
TIME, H
Fig. 23.17 Epoxy resin viscosity compared with time and temperature. (Courtesy of Shell Development Company.)
Materials 511 1000 112" DIAMETER ROUND P R O C E S S E D AT 1 2 " I M I N 112 VLO. SINGLE END ROVING
000
-
600
-
ENDS
47-64
16
x
CLAY
4
0
2 4
TOO LOW,
TOO HIGH,
OCCURS
EXCESSIVE
LOADS ARE
A 2 3
n
400
78
79
00
01
02
03
04
05
GLASS CONTENT (NON-COMBUSTIBLES), 'OF W H I C H A P P R O X I M A T E L Y
86
07
% B Y WT.'
2 X IS R E S I D U A L C L A Y
Fig. 23.18 Epoxy resin pull loads compared with glass content. (Courtesy of Shell Development Company.)
ing part. Table 23.3 lists the target reinforcement volumes for the epoxy system for a variety of reinforcement systems. These values were generated by following the same procedure as for the data in Fig. 23.18. The data shown in Table 23.3 are qualitative rather than quantitative values. They may be used to estimate the required reinforcement volume. Table 23.3 Target fiber volume ranges for epoxy
pultrusion wt.%
All glass roving reinforced composites Multi-end type rovings Single end type rovings
78 77-81
Glass roving and continuous mat reinforced composites
3.175 nun (1/8in) thick cross sections 6.35 nun (1/4 in) thick cross sections
6447 71-74
Carbon fiber reinforced composites All unidirectional tows
67-74
(57-65VOlYO)
23.6.7 DIE TEMPERATURE CONTROL
In polyester pultrusion, there is a wealth of prior experience7 which can be used to provide a temperature set point to produce the desired surface and internal quality of the part with the controller being a thermocouple located some short distance from the entrance to the die. This creates a situation which is independent of the actual exotherm temperature in the curing process. For epoxy pultrusion it is vital that the peak exotherm be understood and controlled. It should not exceed 225°C (437°F) in the hottest region of the part and die. At this temperature homopolymerization will take place within the epoxy resin system and the resin does not need the curing agent to stimulate the cure. The mechanical and physical properties of the structure are degraded under these conditions by the presence of the unused curing agent. For most thin profiles (up to 12.7 mm (0.5 in) thick), a single heating zone is sufficient. The thermocouple should be located in the center
512 Pultrusion CIT
1 2 ' /MINUTE 1 8 ' S T R I P HEATERS
1'-
----
TEYPERATURE ON THE SURFACE OF THE ROD TEYPERATURE I N THE CENTER OF T H E R O D
I
0
6
I
1
I
12 18 24 DIE LENGTH, IN
1
I
I
I
30
36
42
48
Fig. 23.19 Die temperature profile of single zone heating. (Courtesy of Shell Development Company.)
of the strip heater to minimize the overshoot in the epoxy resin system. If the temperature and lag time for the temperature controller. set point is too low, the resin cure rate will For a 12.7 mm (0.5 in) thick cross section the decrease which increases the size of the gel temperature set point of 200°C (392°F) on the zone. At the same time, the rate of shrinkage surface of the die will yield an internal peak and the rate of hydraulic pressure decay is exotherm temperature of 225°C (437°F).Figure reduced and this results in more pressure 23.19 is a graphic example of this single strip within the larger gel zone which increases pull heater profile. Figure 23.20 depicts the graph loads. As the die temperature increases, the of the temperature set point versus pull loads conditions begin to favor reduced pull loads.
700
-
600
-
n 500
-
m A
0 A A 4
3
400 140
150
160
170
180
190
200
DIE CONTROL TEMPERATURE, "C
Fig. 23.20 Pull load compared with die control temperature. (Courtesy of Shell Development Company.)
Materials 513 With section thicknesses up to 12.7 mm (0.5 in) a two zone heating profile can be used and this will probably eliminate the need for radio frequency (RF preheating) to prevent internal cracking. The pull loads are much lower and the temperature decay rate is less than the single zone heating profile. Thus the two zone version will result in a higher degree of cure than the single zone. For thick pultrusions with sections beyond 12.7 mm (0.5 in), the RF preheating method results in faster processing rates without cracking. Figure 23.21 illustrates the die temperature conditions for the same 12.7 mm (0.5 in) diameter rod pultruded with RF preheating. Preheating serves to reduce the temperature differential between the die entrance and the gel zone which results in less volumetric expansion due to temperature and hence less pressure. The temperature lag between the surface and the center of the part is also reduced by RF heating, thus the gel zone is smaller which reduces pulling loads. Increase of the processing rate through the machine will bring the pressure, the gel zone size and the pull loads back to normal conditions. This is how RF heating permits faster processing rates without increasing the pull loads beyond the standard levels.
23.6.8 RESIN MIXING
The best approach to mixing resin system components is to precisely follow the stated recipe for the amount of material to be added and the degree of mixing following the addition. As the ingredients such as catalyst accelerator, pigments, viscosity and extenders, internal mold release and fillers are added, the resin is mixed for a very short time of up to one minute (Table 23.4). The curing agent is left out of the mix until the mix is ready to be added to the pultrusion resin bath. The addition of filler materials requires a high shear mixing and this should be minimized since it generates significant quantities of heat. Just prior to start up, the curing agent should be added and blended at a reduced mixer speed (high shear mixing is only needed for filler addition.)Following start up of the system, make-up replenishment resin batches should be scaled to the depletion rate for the run.If the resin consumption is 7.5 1 (two gallons) per hour, add 7.5 1 (two gallons) per hour. This addition will assist the pot life of the epoxy resin and will dilute it with fresh resin on a frequent basis. If a small bath size can be used, this will increase the dilution effect of the replenishment material. Depending on the size of the bath, this technique can enhance the pot
CIT 200c
Q 18' /MINUTE r
18' STRIP HEATERS
250k
Y
W I-
----
ON THE SURFACE O F THE ROD
-TEMPERATURE
50-
01
I
1
I
0
6
12
1
TEMPERATURE I N THE CENTER OF THE R O D
I
18 24 DIE LENGTH, I N
I
I
I
I
30
36
42
48
Fig. 23.21 Die temperature profile with FF heating. (Courtesyof Shell Development Company.)
514 Pultrusion life of the resin by 800% or more. A large master batch of the resin can be mixed and set aside in advance of the production run.This batch without the curing agent will be stable for up to three days. The curing agent can then be added to a small make-up batch and mixed in just prior to the addition to the system resin bath. A typical formulation used in batch mixing is shown in Table 23.4. Even with the best of conditions in terms of a small resin bath and frequent addition of new batches, the resin mix will ultimately become too viscous for good fiber wetting. A good tip is to provide a large hole in the resin bath with an appropriate plug. This allows a quick drain and a clean and refill with only a momentary pause in the process. With this step, the pot life of the resin bath will be reset to zero. It should also be noted that the plug in the drain hole should not be threaded since a very small amount of resin can cure and lock up the threads. A preferred plug would be hard rubber with a rim, similar to the knock out drain plugs in the floor of an automobile. 23.7 START-UP PROCEDURE
A key factor in trouble free start up is to use the minimum required amount of reinforcement. However, if too much reinforcement is eliminated in start up, sloughing will take place in the die because of insufficient total pressure in the gel zone. Once the sloughing is
initiated there may be potential for significant build up on the surface of the die and this may be difficult to remove. The most troublesome spots in the die are the low pressure, remote areas such as a comer or a small radius. There is a standard process used in pultrusion of purging the die with pure mold release just prior to entry of the resin. Experience and recent tests have shown that the pure mold release is not necessary and may lead to related problems. The normal types and quantities of mold release recommended for use with epoxy were determined by tests at resin suppliers. Any levels of mold release in excess of those recommended will not provide additional benefit to the process. A high concentration of mold release may result in a 'squeeze off' at the die entrance which could work its way back into the resin bath. If a prelubrication step is used, this squeeze off resin quantity must not be allowed to get back into the resin bath. Prior to the resin entering the bath, the die temperature must be stabilized at the set point. For most parts an initial throughput speed of 25.4-30.4 cm (10-12 in) per min is recommended until the cured stock is in the pullers, to minimize loads imposed on the dry fiber. Process rate increase can then be made gradually and for epoxy resins the processing speeds normally will not exceed 45.7 cm (18 in) per min. It would appear that 30.4-35.5 cm (12-14 in) per min provides the best combina-
Table 23.4 Two part batching for long production runs
Part 'A' (in order of addition) [PHR, (wt.%)] 1. Epon@resin 9310 2. Epon Curing Agent@9360 3. Mold Accelerator 837 Wiz Int. 1846 4. Zylac 907 Blend the above for 30 s then add clay 5 . ASP400P Blend clay for no more than 5-10 min
Courtesy: Shell Chemical Co.
Part 'B' 100.00 0.67 0.70 0.40 20.00
Epon Curing AgenP 9360 33 PHR based on resin weight OR Recompute based on the total weight of Part 'A' (33/121.77) 100 = 27.1 and use 27.1 PHR to Part 'A'
Additives 515 tion of processing parameters, pull loads and surface gloss. If RF preheating is used, perform the start up without it at a reduced speed 15.2-20.3 cm (6-8 in) per min to minimize pull loads. When cured stock is through the die and in the pullers, the RF preheater can gradually be started up. The resin temperature entering the die should be monitored and as it reaches about 71°C (160°F)gradually increase the processing speed. Do not recycle the resin which is squeezed off at the die entrance. 23.7.1 TROUBLESHOOTING
The standard problem encountered with epoxy resin pultrusions is poor surface finish or sloughing. The reasons for this have been discussed earlier in this chapter. If the reinforcement level is low then the cure for this is to obviously add some reinforcement. However, if the redorcement level is in accordance with the specifications, then additional reinforcement will increase the pull loads beyond standard. In this case the die temperature profile is probably too low. There is an instrument on the market known as the Gelstar Thermal Analyzer. The thermal analyzer can be used to obtain a temperature profile within the die. From this, the size of the gel zone can be estimated. This is proportional to the lag time between the die and internal temperature profiles. If the temperatures appear to be within limits but the lag time is too large then the processing rate is too fast for the particular cross section within the die. At this point either reduce the process rate or use RF preheating to minimize the problem. These steps will reduce the size of the gel zone. If sloughing is encountered, the part should automatically clean itself up. The typical purge techniques that are common in the industry can be used with epoxy resins. If a portion of the die refuses to clean up, a trick is to insert a copper kitchen ‘Chore Boy’ in the area of the part which has the problem. This will often push the offending plug out or catch onto it and pull it out. The copper gauze will
provide a mild scrubbing action that will clean the surface of the die. 23.7.2 SHUT DOWN
Standard shut down procedures have been developed for polyester resins. The key step is to remove the resin bath or bypass the reinforcement around it. The dry reinforcement should be completely pulled through the die. None of it should be cut out or removed. At this point the die will be free from build up and ready for a restart. The resin drain from the bath should be placed in a container in an area with good ventilation and spill protection. A metal or plastic tray with a surrounding high lip will be satisfactory for spill protection. The resin containers should only be half full. Eventually when it does exotherm, it will become hot, expand and may overflow the container if it is too full. 23.8 ADDITIVES
Inorganic fillers are used to reduce shrinkage at polymerization. They also extend the volume of the resin phase to provide a low cost formulation. These are primary additives. Fillers can frequently constitute the largest proportion of a formulation, second only to the base resin. Fillers are classified according to their particle size, as either coarse fillers or fine. Coarse fillers have an average particle size in excess of 8 pm and are generally the nonfibrous type with low surface area and low oil absorptions. They can be highly loaded into the resin and are easily wetted out by the resin system. Their disadvantage is that they tend to provide poor compound cohesiveness and to introduce localized resin rich pockets and possibly to increase fiber agglomeration during secondary molding processes (as in pull forming).The large particle size filler can be filtered out by a high density roving preform which can lead to large voids in the interior of the pultruded structure. The most common of these fillers are calcium carbonate, aluminum silicate and
516 Pultrusion alumina trihydrate. Silica and coarse talcs are also examples of coarse fillers. Calcium carbonate is primarily used as a volume extender to provide the lowest cost resin formulation where performance is not critical4s. Fine fillers have an average particle size of 5 pm or less and have high surface areas which can produce high viscosities in the formulations. These fillers provide a high order of cohesiveness and will tend to lubricate the pultrusion system. They also help to reduce localized shrinkage owing to their more complete distribution within the polymer. Kaolin clays, hydrous alumina silicate, fine talc, colloidal silica and precipitated calcium carbonate are examples of fine fillers. Clay fillers are used to improve corrosion resistance and where electrical properties are required. They also impart a superior surface finish to the pultruded product. Alumina trihydrate improves flame and smoke generation properties and occurs in applications where consumer or governmental codes are imposed to decrease flammability. Fillers used in the pultrusion process should contain less than 0-5% free water content and should be uniform and free from contamination. Foreign material in the filler may cause localized reaction with off gassing of volatile byproducts and voids or could affect the uniformity of coloring within the pultrusion. Fillers are mixed into the resins in quantities up to 50% of the total resin formulation by weight (100 parts filler per 100 parts resin). Limits of filler addition are based on the viscosity of the system which results from the particle size of the filler and the characteristics of the resin. Wetting agents are sometimes used to add a volume of filler material without increasing formula viscosity. Wetting agents can be added to the filler by the supplier or as an additive during the formulation process for the resin. Air release agents are added in the same manner and will result in more efficient packing by reducing entrapped air in the liquid resin. They also tend to reduce void content in the finished product.
23.8.1 PIGMENTS
Pultruded products are normally associated with bright colors and these colors are normally created by adding suitable colored pigments to the resin system which are then cured into the matrix material. Pigments are generally of three types: (a) dyes; (b) organic pigments; (c) inorganic pigments. These three pigments are characterized by their own individual proper tie^^^. Dyes have good transparency and acceptable brightness. However, they have poor heat resistance and tend to migrate in solution. Organic pigments also have acceptable brightness and brilliance but are not normally as good as the dyes. Weather resistance and W absorption tends to be a problem and the colors may deteriorate and fade after long periods of W exposure. Inorganic pigments are generally the materials of choice. These are usually natural or synthetic metallic oxides, sulfides, or other salts which are heat treated and converted to a dry powder at 600-1100°C (1112-2012°F). Inorganic pigments have superior properties of brightness to those of organic pigments or dyes and are very resistant to weather and migration and have a very high stability under light exposure. The major problem, however, in incorporating these pigment additives into resin formulations are the effects which they have on the cure cycle of the resins. It is possible for them to be involved in the polymerization reaction during the curing of the resins and to become chemically attached into some of the reactive sites. This has a strong effect on the properties of the resulting composite material and will require a change in the temperature/time curing cycle. A large percentage of pultruded composite components are used outdoors and W exposure is a problem. Under these circumstances, titanium dioxide, an excellent UV absorber and whitener, is used as a part of the formulation. Its presence would also naturally create a paler color in the pultruded composite and hence additional quantities of inorganic pigment are
Additives 517 normally required to make the color bright. Zinc sulfide is also a UV absorber which is frequently used in pigment systems. Both titanium dioxide and zinc sulfide seem to have little effect on the ultimate mechanical and physical properties of the pultrusions such as the strength, moduli and impact resistance.
0
0
0 0
The disadvantages of a mechanically fastened joint are:
23.8.2 STRUCTURAL CONNECTIONS
Pultruded FRP composites can be joined using various methods of assembly including mechanical fastening (with plastic or metallic bolts or screws or by doweling with dowels or rivets), mechanically interlocking connections (where molded or laminated inserts lock into the sections being fastened), adhesive bonding, or a combination of these methods50. When structural components are assembled special attention must be given to the rigidity, geometry, fabrication and assembly requirements involved. It has been determined that mechanical fastening and/or mechanically interlocking connections are most suitable for structural connections. Some connections depending upon the geometry and the stresses developed in connection can be strengthened by also adhesive bonding in addition to mechanical fasteners. Each of these systems have their own advantages and disadvantages which are discussed in the following paragraphs.
in preparing the joints, holes must be drilled and sealed, structural members must be trimmed and any required gusset plates must be fabricated; drilling of holes cuts the longitudinal strength fibers in the area of the hole. This can cause high stress levels and stress concentrations; strength of the joint is dependent upon the bearing strength of the composite material; the strength of mechanically fastened joints is also dependent upon the strength of the fastener. Fastener parameters include: 0 0 0 0 0 0
0 0 0
23.8.3 MECHANICAL FASTENERS
If properly designed and fabricated, mechanically fastened connections are the most reliable method of joining pultruded structural sections. In recent years a significant amount of information through empirical testing and prototyping of connections has become available on the bolting and riveting of composites. There are many standard references in this field51-53. The advantages of a mechanically fastened joint are:
surface preparation of the composites is not required; inspection of the joint is relatively easy; the joint can be assembled and connected up to full strength rapidly; ability to disassemble and reassemble.
0 0
clamping force/installation torque limits; washer size for transmission of load; fastener size; hole size and tolerance; joint type; geometry of the fastener layout; composite thickness; rate of loading and the direction; static or dynamic loading; failure criteria; high stress concentrations around fastener.
23.8.4 ADHESIVE JOINTS
Adhesive joints have become popular for the connection of the composite materials since there is not degradation of the composite itself by the bonding process. Adhesives are usually available in solid, paste or liquid form and are classified as either inorganic or organic materials. The majority of the structural adhesives are derived from the organic group and can be
518 Pultrusion
thermosetting resins, thermoplastics, or elastomers. These adhesives are known for their properties of being strong, tough, insoluble and useable over a wide temperature rangeM.The primary advantages of an adhesive joint are: 0 0
0 0
0
0
the properties of the composite material remain intact and are not degraded; adhesives are generally stronger than the composite material being bonded, consequently any failure mode is usually forced into the surrounding primary structural material; good distribution of joint stresses; adhesive bonding can be used to bond dissimilar materials; adhesive joints can be particularly forgiving. Flaws in the joints do not generally degrade the strength of the joint. The main reason for this is that the critical stress location in a bonded joint is usually at the end of the overlap and flaws tend to occur in the center of the joint where the stress level is low; good fatigue and impact loading characteristics.
Disadvantages of the adhesive bonded joints are: 0 0
0 0
0
surface pretreatment such as cleaning and etching must be carefully performed; the preparation and mixing of the adhesive in the correct proportions is critical. Manufacturers instructions must be followed precisely; there is a time limit or shelf life of the adhesive following its preparation; the component parts must be carefully located ;sing holding tools and fixtures and maintained in position with pressure across the joint during the curing of the adhesive; while adhesive joints are normally designed and stressed in shear, there is an additional failure mode owing to peel stress. This is a tensile stress which develops a maximum value at the free end of a
0
double or single lap joint. The stress can be particularly severe in thick, double lapped joints; adhesive bonded joints take time to cure while the resin is setting up and hardening. During the cure cycle, the component parts must be restrained in a fixed position.
23.8.5 MECHANICAL ADHESIVE COMBINATION
Some connections are stronger using a mechanically fastened/adhesive bonded joint. Advantages of mechanically fastened/adhesive bonded joints are: 0 0 0 0 0
higher overall capacities; greater resistance to environmental and thermal deteriorations; less subject to peel stress failure than ‘bonded only’ joint; improved fatigue and impact characteristics; increased joint rigidity.
The disadvantage of mechanically fastened/adhesive bonded joints is that they are labor intensive. 23.9 APPLICATIONS FOR PULTRUDED
PRODUCTS
Applications for pultruded products are many and varied but generally are commercially oriented. The process shows up in the consumer and recreation market, electrical equipment products, corrosion resistance, civil engineering and construction and transportation (automotive, truck, bus and rail). 23.9.1 CONSUMER/RECREATION
The combination of strength, stiffness, fatigue resistance and aesthetic design and coloring makes pultruded products ideal for the consumer recreation market. Applications are: fishing rods, archery bows and arrows, hockey sticks, tent poles, ski poles, playground equipment, fence posts and baseball bats.
Applications for pultruded products 519 23.9.2 ELECTRICAL EQUIPMENT
In this marketplace, strength, electrical insulation and safety are primary attributes of pultruded The following are the most significant applications:fuse holders, ladders, tool handles, electrical conduits5(', cable traysy and power rail covers for subway trains. Figure 23.22 shows a set of pultruded ladders and Fig. 23.23 shows the installation of a power rail cover for a subway train.
Fig.23.23 Third rail cover for rapid transit rail system. (Courtesy of Creative Pultrusions Inc.) 23.9.3 C M L ENGINEERING/CONSTRUCTION MARKET
In this market the properties which are required are strength, modulus, corrosion resistance and nonslip surfaces. Applications are: gratings, stairs, guard rails, bridges and platform^^^, crash barriers, ladder cages, structural supports614, sign posts and signs, light and pedestrian bridges. Figure 23.24 shows an installationof gratings and hand rails in a chemical plant. Figure 23.25 shows a pedestrian bridge and Fig. 23.26 shows the Aberfeldy foot bridge in the UK. This latter applicationis one of the most interesting developments in the application of pultruded sections. This bridge was designed and erected by Maunsell Structural Plastics Division of Fig. 23.22 Pultruded ladders. (Courtesy of Creative Maunsell Engineering in London. The bridge spans the River Tay in Scotland and C O M & S Pultrusions Inc.)
520 Pultrusion
,, Fig. 23.26 Aberfeldy footbridge. (Courtesy of Maunsell Structural Plastics Ltd.) Fig. 23.24 Grating and handrail installation. (Courtesy of North West Fibre Mechanics Ltd.)
two sections of a golf course. The bridge is a double cable stay design and all of the components are made of composite materials with the bridge decking and guard rails being pultruded products. Maunsell has also installed a composite vehicular bridge at Bonds Mill in England which was opened in 1994. 23.9.4 TRANSPORTATION MARKET
Fig. 23.25 Pedestrian bridge in Pennsylvania. (Courtesy of Creative Pultrusions Inc.)
This market is potentially very large. It includes automotive, trucks, busses, light rail, subway trains and passenger trains. The products which are being pultruded for this large array of industries are as follows: drive shafts for trucks66,leaf springs67,68, bumpers, frames and cross members, transportation container bodies, roll up doors, refrigerated truck components, frames for light rail cars, interior structure for passenger trains and subway load carrying beds for small trucks and frangible airport approach masts.
References 521 Figure 23.27 shows a bus interior, fabricated from pultruded sections and Fig. 23.28 displays a frangible airport approach mast. 23.9.5 MISCELLANEOUS
Another market which is using pultruded products is the oil and gas industry for off shore oil well platforms. Application here is for floor gratings, hand rails, stairs and storage buildings and living quarters on the platforms. Another emerging application is in the constant cross section blade for the Darius design of windmills for alternate power. This is just a brief review Of the current aPP1ications. The future of the pultrusion process and its applications is only limited by the scope of human imagination. The market will continuously increase and it is predicted that by the year 2000 the total volume of pultruded products will have tripled over the 1995 levels.
Fig. 23.27 Bus interior showing pultruded composites. (Courtesy of Creative Pultrusions Inc.)
..”
.-
-----c?9.
REFERENCES 1. Goldsworthy, W. Brandt, US Patent, 2 871 911 Apparatus for Producing Elongated Articles from Fiber-reinforced Plastic Material; Issued 2/12/59. 2. Birsa, R. and Taft, P., New Materials Approach for Providing Transverse Strength in Pultruded Shapes, RP/C Reinforced Plastics/Composites ‘84; Composites go to the Market; Papers presented at Technical Sessions of the 39th Ann. Conf., New York. Jan 16-19,1984, Session 1-A, p. 4, 627, SPI Reinforced Plastics/Composites Institute. 3. Taft, P. and Birsa, R., Transverse Strength for Pultruded Parts, Plusf. Engng., 1984,40, (5), 634. 4. 5.
6. Barking: Elsevier Applied Science, 1986,p. 1-46, 012.
Fig. 23.28 Frangible airport approach mast. (Courtesy of Creative Pultrusions Inc.)
522 Pultrusion 7. Martin, J., Pultrusion, Plastics Products Design 19. Beever, W.H. and O’Connor, J.E., Pultruded Thermoplastic Composite Structures, Int. Handbook. Part B. Processes and Design for S A M P E Symp. Proc., 32, 1309,1987. Processes, (ed Miller, E.), New York: Marcel 20. Beever, W.H. and OConnor, J.E., Polyphenylene Dekker Inc., 1983, p. 37-74. Sulphide Pultruded Type Composite Structure, 8. Martin, J. and Sumerak, J.E., Pultruded 42nd Ann. Conf., Plastics/Composites Inst., Composites - The Case Against Aluminum 1987. Extrusions, Pultrusion Technology, Inc., RP/C Reinforced Plastics/Composites ‘84; 21. Wood, AS., Pultrusion is Poised for New Growth and It Won’t be All Thermosets, Mod. Composites go to the Market; Papers presented Plast. Int., 1976, 6(6), 47-9. at Technical Sessions of the 39th Annual Conference, New York. Jan 16-19, 1984, Session 22. Goldsworthy, W. Brandt, Thermoplastic Composites: The New Structurals, Plast. World, 1-D, p. 5, Confer. 627, SPI Reinforced Plastics/ 1984,42(9),56-8. Composites Institute. 9. Laguan, O., Pultrusion: Economic Aspects, 23. Kidd, A.C., Winding and Profile Production Tape, Filament, Pultrusion-Extrusion, Applications and Design, Rev. Plast. Mod., 1985, Reinforced Plastics, In Proc. Electrical Symp., 50(349), 61-6 (Spanish). Bristol, Feb 1975, Paper 5, p. 18 Preprint 627-61 10. Spencer, R.A.P., Developments in Pultrusion, In Developments in GRP Technology - 1, (ed. Harris, 24. Smith, A., Pull Winding Techniques Improve Pultruded Products, Pop. Plast., 1988, 33(4), B.), Barking: Applied Science, 1983, p. 1-36, 42-3. 6272. 11. Martin, J.D., Pultrusion: The Other Process, 25. Shaw Stewart, D.E., Pullwinding Conf. Proc., 2nd Int. Conf. on Automatic Composites, Paper Plast. Engng., 1979,35(3), 53-7. 15, Noordwijkerhout, The Netherlands, 26-28 12. Beck, D.E., New Processes and Prospects in Sept. 1988. Pultrusion, Goldsworthy Engineering Inc., Composite Solutions to Material Challenges: 26. Goldsworthy, W. Brandt, US Patent, 3 674 601 38th Ann. Conf Preprint, Houston, Tex., Augmented Curing of Reinforced Plastic Stock; February 7-11, 1983, Session 6-B, p. 4, Confer. Issued 7/4/72. 627, SPI, Reinforced Plastics/Composites 27. Goldsworthy, W. Brandt, US Patent, 3 793 108 Augmented Curing of Reinforced Plastic Stock; Institute. 13. Roubinet, P., Curved Pultrusion, Composites Issued 2/19/74. Plast. Renf. Fibers. Verre Text, 24(4), July/Aug 28. Parry, T.V. and Wroksky, AS., Effect of 1984,69-73 (French). Hydrostatic Pressure on the Tensile Properties 14. Goldsworthy, W. Brandt, New Technology for of Pultruded CFRP, J. Mater. Sci., 1985, 20(6), 2141-7. Continuous Reinforced Plastics Processing: Its Called ’Pulforming’and It Permits Extrusion of 29. Bibbo, M.A. and Gutowski, T.G.,Analysis of the Pulling Force in Pultrusion, Antec 86. Plastics Variable Cross Section Parts and Curves, Mod. Plast. Int., 1979,9(9),Sept, 100-1. Value Through Technology. Proc. 44th Ann. 15. Goldsworthy, W. Brandt, Pulforming - The Techn. Conf., Boston, April 28-May 1, 1986, p. 1430-2.012. SPE. Changing Shape of Composites. 16. Ewald, G.W., Curved Pulforming - A New 30. Anon, Pultruded Fibre-Reinforcements, Plast. Manufacturing Process for Composite News (Aust.),Nov. 1979, 20. Automobile Springs, Working Together for 31. Hill, J.E., Goan, J.C. and Prescott, R., Properties Strength, 36th Ann. Conf., Washington, D.C., of Pultruded Composite Containing High February 16-20, 1981, Session 16-C, p. 1-6, Modulus Graphite Fibers, S A M P E Qtly, 1973,4, Confer. 012 SPI Reinforced Plastics/Composites (2), 21-7. Institute. 32. Spencer, R.A.P., Advances in Pultrusion of 17. Goldsworthy, W. Brandt, Pulforming Makes Carbon Fibre Composites, Carbon Fibres, 2nd Curved Pultrusions, Brit. Plast. Rubb., Nov 1985, Int. Conf., London, Feb 1974, Paper 21, 140-7, p. 36. Confer. 51FlC. 18. Nepasicky, J and Kannebley, G., Advantages 33. Evans, D.J., Designing with Pultrusions: From and Limitations of Vertical and Horizontal the Idea to the Application, Composite Pultrusion Processing. Examples of Typical Solutions to Material Challenges: 38th Ann. Applications. Conf Preprint, Houston, Tex., February 7-11,
References 523 1983, Session 6-A, p. 5, Confer. 627, SPI Reinforced Plastics/Composites Institute. 34. Browning, J., Synthetic Surface Veils for GRP Laminates, Int. Reinf. Plast. Ind., 1986, 5(5) 14/6. 35. Werner, R.I., Improvements in Means of Evaluating Weathering Characteristics of Pultrusions, Rising to the Challenge: 35th Ann. Conf. New Orleans, LA, Feb 1980, Section 4-E, p. 6, Confer. 627. SPI Reinforced Plastics/ Composites Institute. 36. Heritt, R.W., New, High Performance, Fast Curing Epoxy Resin System for Composites, Composite Solutions to Material Challenges: 38th Ann. Conf. Preprint, Houston, Tex., February 7-11,1983, Session 19-D, p. 7.627, SPI, Reinforced Plastics/Composites Institute. 37. Kershaw, J.A., New Epoxy Resin Systems for Pultrusion, Shell Development Co., Composite Solutions to Material Challenges: 38th Ann. Conf Preprint, Houston, Tex., February 7-11, 1983, Session 6-42, p. 4, 627, SPI Reinforced Plastics/Composites Institute. 38. McQuarrie, T.S., New Generation Resins for Pultrusion, 33rd Ann. Conf., Washington, D.C., Feb 1978, Section %E, p. 5, Confer. 627 SPI Reinforced Plastics/Composites Institute. 39. Howard, R.D. and Sayers, D.R., Development of New Methacrylate Resins for Use in Pultrusion, RP/C Reinforced Plastics/Composites '85. 40 Years of Innovative Technology;Proc. 40th Ann. Conf., Atlanta, GA, January 28-February 1, 1985, Paper 2-A, p. 5, 627, SPI Reinforced Plastics/Composites Institute. 40. Boinot, E and Daspet, Y., Phenolic Resins: A Boon to the Building and Transport Industries, Composites Plast. Renf. Fibres Verre Text, 26, 3, May/June 1986,97-100. 41. Lo Scalzo, E., Phenolic Resins in Advanced Composites, Mat. Plast. Elast., 1988, 3, p. 124-9 (Italian). 42. Hunger, G.A., Pultruding Epoxy Resin, Reprinted from 43rd Ann. Conf. and Focus '88. Proc. 43rd Ann. Conf. SPI Reinforced Plastics/Composites Institute 43. Kiernan, D., Tessier, N. and Schott, N., Modification of Epoxy Resins for Improved Pultrusion Processing, US Army Materials & Mechanics Research Center, RP/C Reinforced Plastics/Composites '85, 40 Years of Innovative Technology; Proc. 40th Ann. Conf., Atlanta, GA, January 28-February 1,1985, Paper 2 4 , p 6 627, SPI Reinforced Plastics/Composites Institute. 44. Sumerak, J.E., Understanding Pultrusion
Process Variables, Mod. Plast., March 1985. 45. Sumerak, J.E. and Martin, J.D., Pultrusion Process Variables and their Effect Upon Manufacturing Capability, RP/C Reinforced Plastics/Composites '84; Composites go to the Market; Papers presented at Technical Sessions of the 39th Ann. Conf., New York. Jan 16-19 1984, Session 1-B, p. 7, Confer. 627, SPI Reinforced Plastics/Composites Institute. 46. Sumerak, J.E., Understanding Pultrusion Process Variables, Mod. Plast., 1985,62(3),58-64. 47. Sumerak, J.E., Understanding Pultrusion Process Variables for the First Time, Conference, Atlanta, GA, January 28-February 1, 1985, Paper No 2-B, p. 8, Confer. 627, SPI Reinforced Plastics/Composites Institute. 48. Armstrong, R.F., Calcium Carbonate, Encyclopedia of Chemical Technology, Vol. 4, 2nd ed., Wiley Interscience, New York, 1964. 49. Anderson, R. and Riddel, R., Effects of Pigments on Pultrusion Physical Properties and Performance, Molded Fiberglass Co.; Morrison Molder Fiberglass Co., Composite Solutions to Material Challenges: 38th Ann. Conf Preprint, Houston, Tex., February 7-11, 1983, Session 6-H, p. 4, Confer. 627, SPI Reinforced Plastics/ Composites Institute. 50. Rufolo, A., Design Manual for Jointing of Glass Reinforced Plastics, US Naval Material Laboratory Report, Navship 250-6341, August 1951. 51. Hodgkinson, J.M., de Beer, D.L. and Matthews, EL., The Strength of Bolted Joints in Kevlar R.P., Proc. Composite Design for Space Application, The Netherlands, (esa SP-243) 15-18, October 1985. 52. Hollaway, L. and Baker, S., The Development of Nodal Joints Suitable for Double Layer Skeletal System made from Fibre/Matrix Composites, Part 7, Paper 21, Proc. 3rd Intern. Conf. on Space Structures (ed. Nooshin, H.), Barking: Elsevier Applied Science, 1984. 53. Green, A.K. and Phillips, L.N., Crimp-Bonded End-Fittings for Use on Pultruded Composite Sections, Composites, 1982,13(3),219-24. 54. Hart-Smith, L.J., Adhesively Bonded Joints for Fibrous Composite Structures, Douglas paper 7740, Long Beach, California, 1986. 55. Anderson, R., Use of Pultruded Reinforced Plastics in Energy Generations and Energy Related Applications, Working Together for Strength, 36th Ann. C o d , Washington, D.C., February 16-20, 1981, Session 22-B, p. 1-3,
524 Pultrusion Confer. 012. SPI Reinforced Plastics/Composites Institute. 56. Morara, F. and Eva, G., GRP Conduits and Poles, Agrosistemi, Macplas, 1985, 10(67), 136-9. 57. Pultrusions for Cable Rack, Brockhouse Group, Eur. Plast. News, 1982, 9(4), p. 36. 58. Mallick, P.K., Qiauw, L.K. and Fesko, D.G., Design and Evaluation of a Pultruded Hybrid Beam, Working Together for Strength, 36th Ann. Conf., Washington, D.C., Febraury 16-20, 1981, Session 1 7 4 , p. 1-5, Confer. 012. SPI Reinforced Plastics/Composites Institute. 59. Head, P.R., GRP Walkway Membranes for Bridge Access and Protection, 13th Reinforced Plastics Congress, 1982, Brighton, November 8-11,1982, Paper 20,97-91, BFP Publn. 293, BPF, Reinforced Plastics Group. 60. Head, P.R., Pultruded Box Beams, Fibreforce Composites, Ltd.; Maunsell Structural Plastics; UK Dept. of Transport; Windfoil Ltd. 61. Anderson, R.A. and Thomas, C., Development of Large Hollow Rectangular Tubes for Structural and Electrical Markets - A Unique Application for Pultrusion, Rising to the Challenge: 25th Ann. Conf., New Orleans, LA, Feb 1980, Section &A, p. 5, Confer. 627. SPI Reinforced Plastics/Composites Institute.
62. Starr, T.F., Structural Applications for Pultruded Profiles, TECHNOLEX, Composite Structures 2; Proc. 2nd Intern. Conf. Composite Structures, Paisley, September 1616,1983, p. 192-216,627. 63. Tickle, J.D., Halliday, G.A., Lazzarou, J. and Riseborough, B., Designing Structures With Pultruded Fibre Glass Reinforced Plastic Structural Profiles as Compared to Standard Steel Profiles, 33rd Ann. Conf., Washington, D.C., Feb 1978, Section SF, p. 8, Confer. 627.SPI Reinforced Plastics/Composites Inst. 64. Owens-Corning Fiberglas Europe SA, Fiberglas in Action: FRP Lighting Poles, Burssels, 1977, Publn, 13-Ch. 4-5, p / 4 12 ins. 16/2/77 6272-6R. 65. Mutch, W., Composite Utility Pole, Plast. World 1987,45, (8), 43. 66. Kliger, H.S., Yates, D.N. and Davis, G.C.R., Driveshafts: The Next Step for Composites?, Aufomot. Engng, 1980,88(3),1OC-3. 67. Roubinet, P. and Delacroix, B., Industrial Development of Composite Leaf Springs, Composites Plast. RenJ Fibres. Verre Text., 26,(3), May/June 1986, p. 79-83 (French). 68. de Goncourt, L. and Sayers, K.H., Composite Spring Systems, Composites Plast., Rent Fibres Verre Text, 1988,28(3), 145-50 (French). 69. BTR Permali RP Ltd, Pultrusion Protects Passengers, Europ. Plast. News, 15(3), 1988,46.
PROCESSING THERMOPLASTIC COMPOSITES
24
James L. Throne
24.1 INTRODUCTION Machinery
Thermoplasticpolymers are seldom converted into products without the time-dependent application of temperature, pressure, shear or other types of mechanical manipulation. The mechanical manipulation of plastics is called 'polymer processing'. Many polymer processes and combinations of polymer processes are used in modern commercial manufacturing. The selection of a process to produce a thermoplastic polymer product from pellets, powder, or other granular forms begins with general characteristicsof the product itself. The two primary concerns to be met in the manufacture of any polymer product are: Will the finished part meet all required, specified and desired design criteria? 0 Can the product be produced at the minimum cost for the projected market size? 0
The first concern focuses primarily on the ability of the polymer to meet mechanical and environmental challenges throughout its functional lifetime. This is shown as the left branch of the Fig. 24.1 schematic'. The second concern deals with the economic ability to process the acceptable polymers into the useful product, and this is shown as the right branch of Fig. 24.1. Commercial polymers are rarely pure. Even 'neat' polymers or polymers that contain no
Handbook of Composites. Edited by S.T. Peters. Published in 1998by Chapman & Hall, London. ISBN 0 412 54020 7
Product Requirements
Shape or P a r t Production Requirement
Electrical Environmental Rigidity Tempera tu r e Possible Processes
Polymer Families
I
F i b e r Type F i b e r Length Other Adducts
ti Concerns
I Compound Grades
I Economics
1
F i n a l Polymer
Fig. 24.1 A schematic for choosing the proper polymer and an attendant process'.
fillers, reinforcements or foam cells, usually have one or more adducts or additives that alter the basic characteristics of the polymer. Table 24.1 gives a short list of some of the adducts used with thermoplastics*. Some of these, such as coupling agents, are vital in achieving the desired final solid mechanical performance of other adducts, such as fillers and reinforcements. Typical fillers used in
526 Processing thermoplastic composites Table 24.1 Adducts in thermoplastic polymers’
Antioxidants Antistatic agents Colorants and pigments Coupling agents Flame retardants Fillers Foaming agents Heat stabilizers Mold release agents Odor suppressors Plasticizers Processing aids Emulsifiers Lubricants Reinforcing fibers Ultraviolet stabilizers Viscosity depressants
thermoplastic polymers are given in Table 24.23. Typical fibrous reinforcements used in thermoplastic polymers are given in Table 24.34. Filled, reinforced and foamed thermoplastics offer great breadth of solid mechanical properties. In many cases, they offer substantial processing challenges, as well. Nearly all thermoplastic processes shape the polymer in its fluid state (The most notable exception to this is thermoforming, where forming occurs when the polymer is in a rubbery state. However, thermoforming depends on the production of sheet that is produced by calendering or extruding the polymer in its fluid state.) Fillers and reinforcing fibers increase the viscosity of the polymer, making it more difficult to shape. As expected, processing difficulty increases with increased filler or reinforcement loading. There are more than twenty major types of polymer processes5.Not all these processes are suitable for thermoplastics and not all thermoplastic processes are suitable for filled, reinforced or foamed thermoplastics. Table 24.4 lists most of the thermoplastic processes that are used with neat, filled or reinforced poIymers. Several of these processes are dis-
cussed in detail shortly. As is apparent, fillers and fibers increase the polymer processing difficulty. In certain instances, as with continuous reinforcing fibers, conventional thermoplastic processing cannot be used. Part geometry is one way of classifying suitable polymer processes, Table 24.56. This classification is further amplified in Table 24.67. Again, not all these processes are suitable for processing all polymers with all combinations of fillers or reinforcements. As noted in Fig. 24.1, the cost of the polymer is only one aspect of the overall economics of product manufacture. Table 24.7 gives a relative comparison of the process costs for filled and reinforced polymers8. Extrusion and injection molding are the primary methods for producing foamed, filled and discontinuous-fiber reinforced thermoplastics. As an example of the growth in molding thermoplastic composites, in the 1950s esentially all injection molded thermoplastics were neat or unfilled and unreinforced. By the early 1990s, filled, foamed and reinforced polymers accounted for,more than 25% (wt) of all injection molded parts. In certain instances, blow molding and rotational molding are possible. Thermoforming or rubbery sheet deformation is now being applied to continuous fiber reinforced polymers. These processes are described below, with the objective of comparing general operating conditions of neat polymers with thermoplastic composites. The technical details of these processes are given elsewhere9-I4.Two axioms apply: 0
0
Axiom I: If the neat polymer is processed in conventional polymer processing equipment, composite versions of that polymer are usually processed in adapted or modified versions of that equipment. Axiom 11: Processing is always more difficult with composite versions of processable neat polymers.
An important corollary also applies:
Rheology,fiber flow and fiber orientation 527 Table 24.2 Fillers for thermoplastic polymers3
Silica products Minerals Sand Quartz Novaculite Tripoli Diatomaceous earth Synthetic amorphous silica Wet process silica Fumed coloidal silica Silica aerogel Silicates Minerals Kaolin (China clay) Mica Nepheline silicate Talc Wollastonite Asbestos Synthetic products Calcium silicate Aluminum silicate Glass Glass flakes Hollow glass spheres Cellular glass nodules Glass granules or cullet Calcium carbonate Chalk Limestone Precipitated calcium carbonate
0
Corollary I: Adding foam, fillers or reinforcements to neat polymers will never improve their processability.
24.2 RHEOLOGY, FIBER FLOW AND FIBER
ORIENTATION
Rheology is the study of polymer flow. Shear flow and elongational flow dominate polymer processing. The great length of polymer chains results in extensive entanglements and complicates the study of neat polymer flow. The
Metallic oxides Zinc oxide Alumina Magnesia Titania Beryllium oxide Aluminum trihydrate Other inorganic compounds Barium sulfate Silicon carbide Molybdenum disulfide Barium ferrite Mica Metal powders Aluminum Bronze Lead Stainless steel zinc Carbon Carbon black channel black Furnace black Ground petroleum coke Pyrolized products Intercalated/exfoliated graphite Cellulosic fillers Wood flour Shell flour Comminuted polymers
economic importance of polymer processing and the technical challenge of predicting molten polymer response to applied load have resulted in an incredibly rich literature15z1. Neat polymer melts are considered as viscoelastic non-Newtonian fluids. Viscosity is the measure of fluid resistance to applied load. The viscosities of oil and water are material constants, independent of shear rate. Fluids of this type are called Newtonian fluids. In steady-state shearing flow, polymers typically exhibit shear-dependent viscosities, as with
528 Processing thermoplastic composites Table 24.3 Fibers for reinforcing thermoplastics
Cellulose fibers a-Cellulose Pulp preforms Cotton flock
differences. Shear rate-dependent viscosity and normal stress differences represent polymer material functions and are not material constants.
Jute
Sisal Rayon Synthetic organic fibers Polyamide (nylon, PA) Polyester (PET) Polyacrylonitrile(PAN) Polyvinyl alcohol (PVOH) Carbon fiber Asbestos fiber Fibrous glass Filaments Chopped strand Reinforcingmat Glass yarn Glass ribbon Whiskers Aluminum oxide (Corundum) Titanium dioxide Boron Boron nitride Boron carbide Metallic fibers Aluminum Stainless steel Copper Tungsten
low-density polyethylene in Fig. 24.2=. When the shearing force on a Newtonian fluid is released, the fluid resistance instantaneously ceases. When the shearing force on a polymeric fluid is released, the fluid exhibits a measure of time-dependent reorganization, the extent of which depends on the extent and duration of the applied forces. Viscoelastic fluids exhibit fading memory of deformation history. This is manifested by normal stress
'p
a
E
v)
l
o
2
m-'
mo Shear Rate, s-'
m1
2 lo2
m3
Fig. 24.2 Shear-rate dependent viscosity of lowdensity polyethyleneat 180°C(356°F) with titanium
dioxide filler in volume YO.(Adapted and redrawn from Ref. 22 by permission of the Academic Press.)
The inclusion of particulates further complicates the rheological behavior of polymers (Fig. 24.2). The effect of filler loading on steady-state shear viscosity of polymers is approximated by:
zw = Y + K j "
(24.1)
where zw is the shear stress at the wall, f is the shear rate, and Y, K and n are empirical constants. This Hershel-Bulkley is the power-law equivalent of the Bingham model for Newtonian plastic fluids. As anticipated, increasing particle surface area to volume increases the viscosity of the polymer, even at the same loading level and particle size distribution (Fig. 24.325). Increasing particulate loading levels usually decreases polymer viscoelasticity as measured by the first normal stress difference (Fig. 24.426). It has been suggested that the shear-rate dependent viscosity of particulate and fiberfilled polymers can be predicted from a
Xheology,fiberflow and fiber orientation 529 Table 24.4 Effect of adduct type on thermoplastic polymer processability
Degree of difficulty (0 = Easy, 9 = Difficult, X = Not done)
Polymer process
Neat
Filled" reinforced
Short-fiber reinforced
Longlfiber organicfiber reinforced
Continuous inorganic fiber reinforced
Continuous
Extrusion Sheet Single screw
0
3
5
8
b
b
5
8
b
b
Extrusion Profile Twin screw
4
5
b
b
Extrusion Compounding Twin screw
5
7
b
b
Extrusion Profile Single screw
Extrusion Foam Pultrusion Injection molding Injection molding Foam Blow molding Thermoforming Compression molding Rotational molding a
2
5
8
X
X
X
X 1 2
X 3
X
X
8
6
8
4
7
X
9 X X
X X
X 9 5 X
2 1 X 2
Short aspect-ratio adducts. Pultrusion replaces conventional extrusion for continuous filament thermoplastic polymers.
temperature-invariant master curve: MFI, log,oMFL, 1
8.86 (T, - Ts) 8.86 (T, - Ts) 101.6 + (T, - Ts) (24.2) 101.6 + (T, - Ts)
where MFI is the melt flow index determined using a melt flow indexer die of L / D = 3.8, according to ASTM D1238. T , is the ASTM recommended test temperature (K), T2 is the
polymer temperature needed for the viscosity determination (K),T, = Tg + 50 (K), and T is the glass transition temperature (K)27. &e experimental results for many polymers and filler and fiber types show good agreement with the master curve. Blow molding and thermoforming are polymer processes that employ melt or rubbery phase stretching. Elongationalviscosity shows an increase in value with increasing filler loading in a manner similar to shear viscosity, Fig.
530 Processing thermoplastic composites Table 24.5 Part geometry as means of classifying processes for filled and reinforced thermoplastic polymers6 Linear forming Extrusion (sheet) Extrusion (profile) Pultrusion Formation of a solid body by injecting into a cavity Unfoamed injection molding Foamed injection molding Formation of a hollow object Blow molding Rotational molding Sheet forming Thermoforming
24.528.As a first approximation, the elongational viscosity for filled polymers is determined in a fashion similar to that for shear viscosities, using equation (24.1)29.For many neat polymers, the elongational viscosity is proportional to three times the shear viscosity. For filled polymers the proportionality is substantially greater than three and is usually shear rate-dependent34 Fiber flow in a shear field is far more complex than particulate flow. Independent fiber closed-orbit rotation as a consequence of shear is dominant in very dilute suspensions.
104 10-2
I
,
I
I
, , , I 1
10-l
,
1
1
1
1
,
100
Shear Rate, s-’
Fig. 24.4 Shear-rate dependent first normal stress difference for polystyrene at 180°C (356°F)with carbon black filler in volume %. (Adapted and redrawn from Ref. 26 by permission of Academic Press.)
As fiber concentration increases, fiber interaction inhibits independent fiber motion. At moderately high fiber concentration or for
?
Fig. 24.3 Shear-rate dependent viscosity of polypropylene at 200°C (392°F) with two types of fillers having particle sizes of 44pm or less. (Adapted and redrawn from Ref. 25 by permission of Academic Press.)
) m3
m-3
m-*
Shear Rate. s 1
1 6
o, , ,
Thermal properties offlled and reinforced thermoplastics 531 fraction, these overshoots can be sustained for several minutes. 24.3 THERMAL PROPERTIES OF FILLED AND REINFORCED THERMOPLASTICS
Most polymer processes depend on energy interchange between the environment and the interior of the melt or solid polymer. Fillers, reinforcing fibers and cells dramatically affect the thermal properties of polymers. Typically, polymers have lower thermal conductivities than inorganic fillers and fibers and higher thermal conductivities than gases that are used as foaming agents. Thermal conductivity, heat capacity, density and thermal diffusivity are the most important. 24.3.1 THERMAL CONDUCTIVITY 005
01
02
03
0 4 05
07
10
Elongation Rate, s”
Fig. 24.5 Elongation-dependent elongational viscosity for polypropylene at 200°C (392°F) with calcium carbonate in volume %. (Adapted and redrawn from Ref. 28 by permission of Academic Press).
Both the solid and fluid thermal conductivities of a filled polymer depend on the relative filler content and the shape of the filler. The Halpin-Tsai equation modified by N i e l ~ e n ~ ~ yields useful values: kc
kP
long fibers, fiber interaction acts to momentarily align fiber segments into bundles. The nature of the polymer flow field then determines whether these momentarily aligned fiber bundles remain oriented as the product is produced. Increased shear implies increased fiber interaction. Fibers not oriented in the flow direction are subjected to increased bending stresses around the fiber bundles. The result is fiber length degradation. In addition to increased normal stress difference with increasing fiber loading, fiber reinforced polymers exhibit substantial shear stress and normal stress difference overshoot at flow inception. These overshoots are attributed to the interactions of fibers moving from an isotropic random state to the more ordered shear flow state31. Depending on flow strength and fiber aspect ratio and volume
-
1+AB$ 1 - BO$
(24.3)
where A = kE - 1, and & is the Einstein coefficient, Table 24.833. $ is the filler volume fraction, P is the maximum packing fraction, Table 24.9%, k, is the thermal conductivity of the composite, kp is the thermal conductivity of the neat polymer, and kf is the thermal conductivity of the filler. B and (T are given as: (24.4) (24.5) The Einstein coefficient, k,, is a measure of the shape of the filler particle. kE = 2.5 for regular shapes such as spheres and becomes large for fibrous or acicular particles. The packing fraction, P, for uniformly sized particles varies in value from 0.52 for random packing to 0.91 for
532 Processing thermoplastic composites
Table 24.6 Classification of polymer processes by size and shape for filled and reinforced thermoplastic polymers7 _______
Processes
2
2
3
Blow molding
Hollow,
Platen
4
5
6
x
x
7
8
9
2 0 2 1
X
X
thin Calendering
wall Sheet
Compression molding Sheet extrusion
Sheet
Profile extrusion Injection molding Injection molding (foam) Pultrusion Rotational Molding Thermoforming
Linear
Linear Hollow
Width of roll, Platen Width of roll, Die Die Platen Platen Die
Thin wall Platen
X
X X
X X
X X X X
x
x
x
x x
x x
x x
x x
X
x
x
x
x
X
X
X
X
X X
x x
X
1 = Shape limitation 2 = Factor limiting maximum size 3 = Complex shapes 4 = Controlled wall thickness 5 = Open hollow shapes 6 = Closed hollow shapes 7 = Very small items 8 = Plane area greater than 1m2 9 = Inserts 10 = Molded-in holes 11 = Threads
hexagonal packing. For fillers with random sizes, P is typically about 0.85 to 0.9. Thermal conductivity is a tensor quantity, with unique values in each of the three principal directions. Because filled polymers are Table 24.7 Ranking of polymer processes according relatively isotropic, the principal values of thermal conductivity are usually equal. For to unit costs continuous carbon graphite filament composites, the thermal conductivity in the fiber Process cost direction is usually much greater than that in Calendering the transverse directions. The packing fracInjection molding Very low tions are determined from Table 24.9 for the Blow molding fiber and crossfiber directions. For random Foam injection molding mats, the typical packing fraction value is Profile extrusion Low about 0.5 in the cross-fiber direction. Rotational molding Sheet extrusion Thermoforming Compression molding Pultrusion Machining
24.3.2 HEAT CAPACITY
Medium High
The polymeric heat capacity or specific heat is the isobaric change of enthalpy with temperature:
Thermal properties of filled and reinforced thermoplastics
533
Table 24.8 Values of the Einstein coefficient, k,, for various types of fillers33
Filler type
k,"
Spheres, one size, maximum packing Spheres, random close packing Spheres, random loose placing Rods or ellipsoids, random packingb Aspect ratio = 2 =4 =6 = 10 Mixed sizes, irregular shapes, minimum surface area Mixed sizes, plates flakesc Agglomerates of spheresd Agglomerates, generally'. a
e
2.50 2.50 2.50 2.58 3.08 3.80 5.93 4.00 5+ 2.5/$,
kd40
Correction factors for the mechanical case of Poisson's ratio, Y , of the matrix: V Factor 0.50 1.00 0.40 0.90 0.35 0.87 0.30 0.84 0.20 0.80 At high rates of shear and low @a, for the rheological case, these shapes tend to orient and to reduce &. These values may be approximated from oil absorption data or more precisely determined by intrinsic viscosity. @, is the volume fraction of the agglomerate that is spherical. @ is the volume fraction of particles in the agglomerate and k, is the appropriate value for the particle shape. These values are generally less than the maximum packing fraction for the particular particle shape.
Table 24.9 Effect of filler particle shape and packing type on maximum packing fraction, P, for uniformly sized particlesM
Particle shape
Packing type
Maximum packing fraction, P
Sphere Sphere Sphere Sphere Sphere Sphere
Hexagonal close packing Face-centered cubic Body-centered cubic Simple cubic Random close packing Random loose packing
0.7405 0.7405 0.60 0.5236 0.637 0.601
Fiber Fiber Fiber Fiber
Parallel hexagonal packing Parallel cubic packing Parallel random packing Random orientation
0.907 0.785
=
(%)
(24.6)
P
The specific heats of fillers and reinforcements
0.82
0.52
are constant. The enthalpies of most amorphous polymers such as polystyrenics, acrylics and polyimides are nearly linearly dependent on temperature and so their specific heats are
534 Processing thermoplastic composites nearly constant. Semicrystalline polymers such as polyolefins, liquid crystal polymers and polyketones have enthalpies that exhibit the effects of melting and so their specific heats are temperature-dependent. The specific heat of a mixture of materials is given as: (24.7)
adduct loading. Thermal diffusivity also increases with increasing void fraction in cellular composite^^^. As a result, the rates of heat addition and of heat removal increase with increasing adduct loading. 24.4 EXTRUSION
Extrusion is a means of producing continuous, linear products in a steady state fashion. Both single-screw and twin-screw extrusion processes are used to produce sheet and profiles from composite thermoplastics. 24.3.3 COMPOSITE DENSITY Single-screw extruder length-to-diameter The density of a mixture of materials is given ( L / D ) ratios range from 15:l to 48:l with 18:l as: to 30:l being usual. Twin-screw extruder L / D s range from 12:l to 30:l. Figure 24.6 is a (24.8) schematic of a typical singlescrew extrudes6. The process is generally characterized as havwhere pc is the density of the mixture, xi is the ing four sequential segments37,38: weight fraction and pi is the density of the ith 0 Solids conveying, where the polymer pelspecies. lets, flake, powder or granules are conveyed from the extruder hopper to the conveying 24.3.4 THERMAL DIFFUSMTY flights of the extruder. The particulates are compacted and begin to heat by shearing Thermal conductivity is the key material propcontact with the metal screw and barrel. erty in steady state heat transfer. The heat flux, Table 24.10 tabulates the polymer character9, is proportional to the thermal driving force istics that are important during solids as: conveying. Methods of determining some dT of these properties are discussed in Sections 9 =k(24.9) dx 24.2 and 24.3. 0 Plasticating or melting, where the gap where T is temperature and x is distance between the barrel and the root of the screw through the composite. is reduced, the polymer melts and the comThe proportionality is thermal conductivity. pressed solid cake breaks up. The degree of In transient heat transfer, thermal diffusivity is compression usually depends on the morthe proportionality, as: phological nature of the neat polymer. Crystalline polymers need greater compres(24.10) sion ratios than amorphous polymers. Polymers with particulate adducts need a where t is time and a is thermal diffusivity, lower compression ratio than neat polygiven as: mers. Table 24.10 also includes the polymer properties that are important in plasticaa = - -k- (24.11) tion. PCP 0 Melt pumping, where the fully fluid For most filled and reinforced polymers, therpolymer is pressure- and temperaturemal diffusivity increases with increasing where xi is the weight fraction and cp, is the specific heat of the ith species.
Extrusion 535 Table 24.10 Polymer properties important in extrusion Solids conveying
Frictional coefficients of particulate polymers with various metallic surfaces Temperature-dependent thermal properties of particulate polymers: Thermal diffusivity Thermal conductivity Heat capacity Bulk density Polymer density Temperature dependent modulus of polymer Temperature dependent yield strength of polymer Plasticating
Temperature dependent and shear rate dependent melt viscosity, Density of polymer melt Density of polymer cake in solid bed Polymer melt thermal conductivity Melt pumping
Melt viscosity dependency on Temperature Shear rate Pressure Polymer melt thermal conductivity Viscous dissipation Thermal and shear degradation potential Extruder die
Polymer shear sensitivity Degradation potential - temperature limitation Melt fracture potential Extrudate swell Rheological characteristics Normal stress difference Temperature dependent elongational viscosity
0
conditioned for the extruder die. The primary polymer property is shear viscosity. Other important properties are given in Table 24.10. Extruder die, where the polymer melt is shaped a n d presented to the take-up equipment39. The extruder die shape depends on the product being produced. For example:
if the extruder die is annular, the resulting product is a hollow pipe or tube. This tube or pipe is also a parison for extrusion blow molding; - if the extruder die is irregular, the resulting product is called a profile. Some of the major polymer properties that are important in profile extrusion are given in Table 24.10.
- if the extruder die is slot-like, the result-
Extrusion dies are dissipative. That is, the polymer exhibits pressure drop through the die equal to that provided at the tip of the extruder screw.
ing product is a planar sheet; - if the extruder die is cylindrical, the
resulting product is a rod;
-
536 Processing thermoplastic composites
\ I
I
H o m b F e e d Throat
Thermocouple Well
Barrel
'
Heate; Band
Main Thrust Bearing
Q Motor
Fig. 24.6 Schematic of conventional single screw extruder without extrusion die. (Redrawn from Ref. 36 by permission of Carl Hanser Verlag.)
Twin-screw extruders are used extensively in producing linear composite products. Figure 24.7 is an illustration of a cylindrical intermeshing twin-screw d e ~ i g n ~ ~ lTwin-screw ~l. extruders are classified according to the relative screw rotational directions, whether the screws intermesh and the relative screw speeds, Table 24.11. Comparative twin-screw performance is given in Table 24.12. Despite some important shortcomings, twin-screw extruders are desired for low shear and controlled feed rates, important aspects of compounding composite thermoplastics as well as extruding them into uniformly consistent products. Increasing time in the shear field results in a linear decrease in fiber length (Fig. 24.847).
Fiber length degradation is much less influenced by screw speed than mixing time. As a result, twin-screw compounding extruders provide less fiber length degradation than do tandem compounding single screw extruders. In fiber-reinforced extruded products, the average fiber orientation for discontinuous fibers is up to about 20" from the axis4s.The orientation is the result of converging flow from the extruder screw tip to the die end. (Converging flow is one of the standard die design critia for neat polymer extrusion. Accelerating flow allows molecuIar alignment in the axis direction and tends to minimize extrudate swell shear d e ~ e n d e n c y ~ ~ . ) Extrudate swell decreases with increasing
Cornpresnion Precompression Metering
I-
Gas-Meit Mixing I_ \ I I
Feed
Preheating
L I
&
L
Fig. 24.7 Schematic of standard configuration of cylindrical twin screws40.(Redrawn by permission of Carl Hanser Verlag.)
Extrusion 537
I
10
0
I
1 10
20
33
I 40
TiO, Volume Percent 96
Post-Injection Molding
0
05
10
15
m
2.0
e '
Fig. 24.9 Effect of volume fraction of titanium dioxide on extrudate swell ratio for high-density polyethylene at 180°C (356°F). Capillary L / D = 28 and shear rate = 10 s-l. (Redrawn from Ref. 50 by permission of Academic Press.)
25
Mixing Length, Diameters
Fig. 24.8 Effect of residence time in shear field on glass fiber length in injection molding and extru~ion~ (By ~ . permission of Society of Plastics Engineers.)
Table 24.11 Classification of twin-screw extruders Intermeshing screws' Corotating screws Low speed extrusion for profiles, foams, filled polymers, short-fiber profiles High speed extrusion for compounding and devolatilization Counter-rotatingscrews Conical extrusion for profiles Cylindrical extrusion for profiles Non-intermeshing screwsb Counter-rotating separated screws With blades for kneading, compounding High speed for in situ polymerization Corotating screws Not used in practice Counter-rotating tangential screws High speed for compounding, devolatilizing Low speed for plasticating fluffy, bulky regrind a
Intermeshing screws are also classified as h l l y or closely intermeshing or partially intermeshinp Non-intermeshing screws are also classified as separated non-intermeshing and tangential non-intermeshing screws4z.
538 Processing thermoplastic composites filler loading (Fig. 24.950)and with increasing fiber orientation in the hoop direction. Orientation in the extrusion direction is desired for profiles that are designed for strength in the bending direction. For pipe and tubing, on the other hand, the strength in the hoop direction is half that in the axial direction for an isotropic polymer. Reinforcement and hence fiber orientation is desired in the hoop or cross-extrusion direction. This is achieved by using a diverging die section following the converging section (Fig. 24.1O5I).The amount
of hoop reinforcement is related to the channel width expansion52. 24.5 INJECTION MOLDING FILLED AND REINFORCED THERMOPLASTICS
Injection molding is a means of producing discrete products on a cyclic b a ~ i s ~The ~,~~. injection molding machine consists of two major parts: the plasticating and pumping section, and the clamping mechanism.
*-
24.5.1 THE PLASTICATING AND PUMPING SECTION, FIG. 24.1155
Annular Die
Fig. 24.10 Expanding mandrel extrusion die with converging section for extrusion of highly fiberreinforced polymers. (Redrawn from Ref. 50 by permission of Society of Plastics Engineers.)
Hopper
A single Archimedean screw similar to an extrusion screw acts to convey and compress the solid thermoplastic, plasticate and melt the polymer and melt convey or pump the polymer melt through a non-return valve or check ring to an accumulation region ahead of the screw. As the polymer accumulates between the screw tip and the nozzle, it pushes the screw backward away from the nozzle, effectively shortening it. When a suitable amount of polymer is accumulated, the mold is closed, the nozzle abuts the sprue, the screw advances and the melt is pushed into the mold. Injection molding screws have L / D ratios of 15:l to 30:l.
f r l \
Hydraulic Piston
Nozzle
Plasticating Screw
W
Non-Return Valve Frame
\ Electromechanical Drive
Fig. 24.11 Schematic of plasticating and injection portion of conventional reciprocating screw injection molding machine. The nozzle inserts into the sprue of the mold mounted on the press shown in Fig. 24.12. (Redrawn from Ref. 55 by permission of Carl Hanser Verlag.)
Injection molding filled and reinforced thermoplastics 539 Table 24.12 Advantages and disadvantages of twin-screw extrudersM Advantages Controlled compaction of powders in feed zone Powder feeding independent of friction with screw or barrel Starved feed decouples feed rate, screw speed and extent of viscous shear heating Rapid but gentle heating and plasticating of thermally sensitive polymers Kneading action provides superior thermal and melt mixing and homogenization Control of pressure build-up by proper element selection High outputs at low speeds, minimizing shear heating Ability to custom design processing sections Gas injection easy to locate Addition of adducts, fillers, reinforcing elements relatively straight forward Comparatively little wear on extruder elements when processing aggressive fibers, fillers Comparatively little fiber attrition Disadvantages Equipment cost per unit output very high Output limited when compared with tandem extruders Screw wear harder to predict More difficult to mix gases into polymer melts in direct gas injection Maintenance tends to be specialized, expensive Although flexibility in changing mixing elements touted, no real way of determining a priori what mixing elements are best for optimum throughput" Pressure build-up entering the die is less effective The thrust bearing remains the primary mechanical weakness a
Recently, computer models have been developed to aid in understanding polymer flow in certain elements such as forward pumping screw elements, backward pumping screw elements and kneading disc elements45.Owing to the complex geometry, twin-screw extruder configuration and element design has not achieved the sophistication of single screws46.
The length ratio of solids conveying to plasticating to melt pumping is 50:25:25. Injection molding screw compression ratios are usually the same as those for extrusion. 24.5.2 THE CLAMPING SECTION
Figure 24.1256shows a fully hydraulic clamp, one method of holding the mold halves closed against the pressure of the injecting melt. The clamp consists of a stationary platen and a moveable platen. The platens are guided open and closed along tie bars. The polymer is introduced through the sprue, a hole in the stationary platen. The moveable platen contains a means of applying pressure to hold the mold halves closed. Hydraulic and mechanical
methods are common means although new electric drives are considered to need less maintenance. In addition, means for ejecting the part from the mold cavity are usually attached to the moveable platen. The polymer is transferred from the accumulator section of the screw into the mold cavity by ramadvancement of the screw at relatively high shear rates of 100-10000 s-l, with transfer times of seconds. Since polymers are comp r e ~ s i b l eat~ ~injection pressures and exhibit decreasing density with increasing temperapressure is applied to the polymer in the mold cavity, runner system and transfer lines after the cavity is filled, until the polymer cools sufficiently to hydraulically seal the mold cavity. This packing pressure is most important in
540 Processing thermoplastic composites Tailstock Platen
Oil Inlet for Closin
Oil Inlet for Opening
L Traversing Cylinder
LMoving
Platen
Stationary Platen
Fig. 24.12 Schematic of hydraulic clamping portion of conventional injection molding machine. (Redrawn from Ref. 56 by permission of Carl Hanser Verlag.)
Table 24.13 Polymer properties important in injection molding Particulate polymer frictional coefficients Solid and liquid polymer thermal properties Thermal diffusivity Thermal conductivity Heat capacity Particulate bulk density Pressure and temperature dependent melt density Polymer shear sensitivity Degradation potential - temperature limitation Shear and temperature dependent viscosity over the shear rate range of 0-10 000 s-I Melt fracture potential
Extrudate swell Rheological characteristics Normal stress difference Temperature dependent elongational viscosity Melt compressibility Volume expansivity Isothermal compressibility Pressure dependent shrinkage Crystallization kinetics Heat of crystallization Rate of crystallization
Injection moldingfilled and reinforced thermoplastics 541 the injection molding of all polymers, including heavily filled and reinforced compounds. A tabulation of polymer properties important in injection molding is given as Table 24.13. Note that many of these properties are important in extrusion, as well. For filled and reinforced polymers, mold design is critical. There are several elements to a mold:
result exhibit differential shrinkage values that increase with increasing loading (Fig. 24.13j9). As a first approximation, the flow of a filled molten thermoplastic can be considered similar to the flow of the neat polymer, regardless
Cross-Flow
Sprue and runner system. The sprue directs the polymer from the injection molding machine nozzle into the mold body through the stationary platen wall. The runner system directs the molten polymer to the appropriate cavities. The gate. Polymer flows into a mold cavity through a constriction called a gate. The size and location of the gate are critical to injection mold part performance, as noted below. Mold temperature control. Coolant lines are normally placed parallel to the machine platens. Adequate coolant flow to all part surfaces is important in minimizing part distortion and warpage. Part removal system. Typically, parts are molded under substantial pressure of 10-15 MPa. Ejector pins and rings are used to press the part from the mold surface after the mold has opened. In addition, other devices such as sliding cores and unscrewing devices are employed to meet certain design criteria.
?
Crystalline neat polymers exhibit greater overall shrinkage and more differential shrinkage than amorphous neat polymers. Filled polymers exhibit lower overall shrinkage but may show substantially greater differential shrinkage, called warp or cupping, than neat polymers. Typically, the value for shrinkage of a filled or reinforced polymer is less than that for the neat state of the polymer, and the value shows substantially less injection-pressure sensitivity. Fiber-reinforced polymers exhibit reduced in-flow shrinkage and greater crossflow shrinkage than neat polymers and as a
;
Diameter
Warp = CuplDiameter I
I
-
Cross-Flow I
Relative Measure of Warp
0.5 W 4-
10
20
40
Glass Fiber Content, %(wt) Fig. 24.13 Differential shrinkage, cupping and warping of glass fiber reinforced polyacetal homopolymer (polyoxymethylene).(Redrawn from Ref. 59 by permission of Carl Hanser Verlag.)
542 Processing thermoplastic composites of the loading level. The amount of force needed to transfer the polymer from the injection molding machine to the mold cavity increases with increase in apparent shear viscosity. Since the transfer pressure is usually fixed by the machine hydraulic system, the transfer rate usually decreases with increasing filler loading. Usually, sprue, runner and gate dimensions are increased to accommodate the lower compressibility of the filled polymer melt. Solid particles tend to migrate away from planes of highest shear. Since shear is greatest near the mold confining surfaces, parts of highly filled polymers tend to have resin-rich surface layers. These layers are typically 10-100 p in thickness. Similarly, weld lines or planes perpendicular to meeting flow fronts, also tend to be resin-rich and are usually weaker than regions on either side of the weld line60. Increasing levels of filler tend to minimize the problem of jetting, where a neat polymer forms a thread rather than a radial disk as it enters a mold cavity from the gate. For fibrous reinforcements, where the L I D aspect ratio is greater than about lO:l, orientation during flow is very important. The fiber rotation attempted in the shear field is complicated by the fountain-like flow of the polymeric mass as it advances through the
mold system (Fig. 24.1463).(It has been mathematically shown that the fountain flow effect holds for cylindrical and planar flow field&. Even at very high fiber loading, individual fibers or fiber aggregates align themselves with the flow streamlines6z.)As is apparent from Fig. 24.15@,there is no appreciable shear field in the center core. As a result, fiber
-wall
Shear Rate
L-
Centerlune
wan
Fig. 24.15 Typical shear rate and velocity profiles for nonisothennal flow during injection mold filling of a cavity. (Redrawn from Ref. 64 by permission of Society of Plastics Engineers.)
Fiber Bundle Orientation
Advancing Flow-Front
Flow Streamlines
Cdd Wall
Fig. 24.14 Fountain flow in injection molding, showing fiber bundle behavior along a streamline and development of frozen layer at cold walls. (Redrawn and reinterpreted from Ref. 63 by permission of Society of Plastics Engineers.)
Injection moldingfilled and reinforced thermoplastics 543 bundles move preferentially from the upstream organizing region transverse to the center axis toward the flow front. In doing so, they orient as the advancing front is formed. As shown in the schematic, the interface between the molten core and the frozen skin grows rapidly into the melt once the advancing front has passed. The oriented fibers are aligned parallel to the flow direction. As with extrusion, the surface at the mold surface is resin-rich and, in many cases, essentally fiberfree. When the mold is full, the structure shows three, five, seven and even nine layers in cross section (Fig. 24.16'j5,@j): the resin-rich surface layer, up to 10 mm (0.4 in) in thickness; a relatively thin layer of fibers oriented in the flow direction, with some of these fibers tipped toward the flow axis; a relatively thick layer of fibers oriented transverse to the flow direction, with a substantial amount of in-plane randomness, depending on the nature of the part being
0
molded; at the very center, a very thin plane of fibers that are oriented in the flow direction.
The general nature of flow into an injection mold cavity through a constriction or gate is radial. Since there is substantial in-plane stretching, fiber orientation in the gate region of long narrow molded parts are quite similar to fiber orientation in center-gated axisymmetric molded parts. If there is no in-plane orientation, cores typically have random planar orientation. Typically, the thickness of the transverse fiber orientation in the core decreases with distance from the gate. For certain polymer-fiber combinations and certain injection speeds and mold temperature, the resin-rich surface layer may not be apparent. In fact, fiber prominence on molded part surfaces is apparent at long flow length-to-cavity thickness ratios, in regions very near the gate when the flow is turned as it enters the mold cavity, at low neat polymer viscosities, for cold polymer melts, and at slow injection rates.
Flow Direction
Transverse Orientation Inflow Orlentation Fiber-Free Layer
In-Flow Orientation Transverse Orientation
Fig. 24.16 Schematic of the development of fiber orientation in injection molding,showing disappearance of fiber-free surface layer and centerline transverse fiber orientation as flow proceeds into the mold cavity. (Redrawn and reinterpreted from Ref. 65 by permission of Society of Plastics Engineers.)
544 Processing thermoplastic composites
Furthermore, the center line or plane of in-axis fibers is usually the result of continuing injection as the flow channel freezes closed. The resulting high shear orients the fibers in the flow direction. This layer may not be apparent in all fiber-reinforced parts. In very thin parts, fibers are nearly always oriented parallel to the mold surface67.The flow in an injection mold is mathematically modeled using Hele-Shaw or creeping flow6*,69.
-
Weld lines are particularly bothersome with fiber-reinforced polymers. The primary concern is the undesirable fiber orientation at the interface between two advancing flow fronts (Fig. 24.17”O). Trpically the reZative weld line strength decreases with increasing filler loading and aspect ratio71(Table 24.14).
-
Flow Direction
Flow Direction
Weld Line
Weld Line
Fig. 24.17 Fiber orientation at the interface of two impinging flow fronts, creating a weak weld line. (Redrawn from Ref. 70 by permission of Carl Hanser Verlag.)
Table 24.14 Relative weld line strength for neat and reinforced polymer^^*,'^ ~~~
Polymer
Tensile strength retention for various glass loadings, YO 0
Polypropylene SAN Polycarbonate Polysulfone Polyphenylene sulfide Nylon 66 (PA-66)
86 80 99 100 83 83-100
10
86 38 87-93
20
30
47
34 40 64 62
40
20 56-64
Thermoforming and compression molding 545 24.6 THERMOFORMING AND
COMPRESSION MOLDING
For neat polymers, thermoforming and compression molding are different disciplines74,75. Thermoforming begins with a formed sheet of plastic that is heated to the rubbery state of the polymer, usually a few degrees above its glass transition or melting temperature. The rubbery sheet is then pressed with relatively little differential force into or onto a cooled singlesided mold and held there until the polymer temperature is substantially below the forming temperature. (In traditional vacuum forming, the space between the sheet and the mold surface is evacuated, thus applying differential pressure of up to 0.1 MPa. In pressure forming, air pressure is applied to the free surface of the sheet, thus applying differential air pressure of up to 1Ml'a76.) The desired part is then trimmed from the web. Compression molding is usually reserved for thermosetting polymers, but certain thermoplastic polymers such as UHMWPE, PTFE, and certain polyimides are compression molded. These polymers are characterized as having very high viscosities even at temperatures hundreds of degrees above their melting or processing temperature^^^. A polymer of this type is compressed as a powder into a preform, heated in a convection oven to the processing temperature, and transferred to a compression molding press where it is compressed between heated mold halves. Typical molding pressures are 7-35 MPa. The mold is then cooled until the formed part temperature is substantially below the polymer processing t e m p e r a t ~ r eTable ~ ~ . 24.15 gives polymer properties important in thermoforming. For filled, foamed and reinforced polymers, the boundaries between thermoforming and compression molding blur. Frequently, the composite sheet forming process is simply called stamping79.Fillers and discontinuous fibers stiffen the polymer so that even at the upper forming temperature of the polymer, substantial differential force is required to
Table 24.15 Polymer properties important in thermof orming Temperature dependent polymer hot strength Elastic modulus at forming temperature Elongational viscosity at the forming temperature Strain-rate hardening at high elongation Other properties that are important in extruding polymer sheet Temperature dependent thermal properties of rubbery solid polymer: Thermal diffusivity Thermal conductivity Heat capacity Polymer density
form the composite from the planar state to a useful product. Mechanical means such as matched dies, or hydraulic forces using superplastic aluminum or polyimide films, replace pneumatic forces when the differential forming pressures exceed 1 MPa. The foamed polymer, on the other hand, cannot be heated to the same forming temperature as the unfoamed polymer without dramatic cell collapse. As a result, foamed polymers are formed at temperatures substantially below forming temperatures for the unfoamed polymers. Mechanical forces, such as matched dies, are used for forming foamed polymer sheet into useful products. Typically, the ratio of polymer modulus at forming temperature to applied pressure, E ( T ) / P , should be in the range of 2-10 with a value of 5 most typical for traditional thermoformingsO.As an example, the temperature-dependent flexural modulus of glass fiber-reinforced polyetherimide is shown in Fig. 24.W. At 200°C (392"F), the modulus of neat PEI is 2 MPa. The pressure required to thermoform this polymer at this temperature is about 0.4 MPa. For 30% (wt) glass fiber-reinforced PEI the modulus is 7MPa and the required pressure is about 1.4 MPa. Pneumatic pressures at this level are possible, but mechanical forming is preferred. In many cases, the elastic modulus of the composite exibits similar temperature dependency to the elastic modulus of the polymer matrix,
546 Processing thermoplastic composites 10,
I 0
I
I
I
I 50
I
I
103
150
Temperature, "c xa
0
zm Temperature. "F
3M
with the isothermal value of the modulus increasing monotonically with increasing filler or fiber loading (Fig. 24.19a2). Thermoforming is a surface-generating process. That is, the total area of the finished part plus web is greater than the total area of the initial sheet. The dominant method of generating surface is biaxial stretching of the rubbery solid polymer. The shape of the temperature-dependent tensile stress-strain curve
0
Fig. 24.18 Temperature dependent flexural modulus of glass fiber reinforced Dolvetherimide. (Redrawn from Rei. s'l by permission of Carl Hanser Verlag.)
of the rubbery sheet is key to the forming process. Figure 24.20 is a schematic of a temperature-dependent stressstrain curve for a neat polymer that exhibits a yield at low temperat u r e ~ The ~ ~ . effect of filler and discontinuous fiber on the shape of this curve is shown in Fig. 24.21M.Typically, the initial slope of the curve, the tensile modulus increases, the yield point disappears and the elongation at break decreases rapidly with increasing fiber or filler loading. This implies that the forming of composite thermoplastics requires high temperatures, substantial forces and the parts
oo 10
20
33
40
Glass Fiber Content. %(wt)
Fig. 24.19 The effect of glass fiber loading on flexural modulus of polysulfone at 25°C (77'F). (Redrawn from Ref. 82 by permission of Carl Hanser Verlag.)
Elongation
Fig. 24.20 Schematic of temperature dependent
stress-strain curve for a neat polymer exhibiting a yield point.
Thermoforming and compression molding 547
Elongation
Fig. 24.21 Schematic of the effect of fiber content on isothermal stress-strain curve for a polymer exhibiting a yield point when unreinforced.
so produced are restricted to relatively shallow draws. Typically, polymers containing nonwoven continuous fiber mat and short- and long-fiber chopped fiber mat are formable with matched metal dies. Composites containing woven fiber mat are not as formablea5.For very long glass fiber- and continuous fiber-reinforced composites, the extensibility of the sheet is so restricted that even shallow draw parts cannot be formed without substantial polymer migration (the squeezing of the polymer matrix from the fiber bundle is called ‘percolation’), compression buckling, pleating wrinkling and fiber breakage when standard matched die molding techniques are used. (Note that this problem is not unique to fiber-reinforced thermoplastic composites. Paper, organic synthetic paper and mixed fiber paper are nonwoven fiber structures that are very difficult to form into deeply drawn product^^,^^.) It is well known that buckling, folding and pleating are minimized by keeping the sheet under tension throughout the forming process. Further, for composites with limited extensibilities, the material that makes up the formed shape must come from the region outside the formed
n
n
Drive Shaft
Cartrldge Heater Bottom Mold Half
Bottom Platen
Mold Open
Mold Closed
Fig. 24.22 Schematic of a mold designed to slip-form continuous fiber-reinforced composite. (Redrawn from Ref. 88 by permission of Society of Plastics Engineers.)
548 Processing thermoplastic composites lntraply Shear
Resin Percolation Through Fiber Bundle
Transverse Squeezing Flow
F
I,
IP
Fig. 24.23 Schematic of fiber-resin matrix interaction during deformation for several types of deformations. (Redrawnand reinterpreted from Ref. 93 by permission of Society of Plastics Engineers.)
shape. Slip forming is the principal method for accomplishingthis (Fig. 24.Zm90).The key is to supply sufficient tension to the sheet to minimize folding but not an excess amount that will tear the very hot ~ h e e t ~ l , ~ ~ . Bulk mechanical deformation of the reinforced sheet is not the only concern when forming reinforced sheet93.Figure 24.23 shows several fiber-matrix interactions that take place locally during composite thermoforming94.Intraply shear, interlaminar shear or slip and interlaminar rotation are local shearing effects that involve the fiber. Local resin flow between fibers, fiber bundles and plies allow local distortion of the matrix. Matrix percolation and squeeze flow are predominantly resin effects. Voids and delaminations are microscopic defects that are attributed to local fiber-matrix interactions during forming. Compression molding is flow molding or ‘squeezingflow’ of polymer between two mold halves (Fig. 24.2495).The flow behavior for a filled or short-fiber reinforced polymer mimics that for the neat The flow is usually characterized as planar radial with the wavefront being fountain-like as with injection molding. Fiber orientation is quite similar to that observed with center-gated injection mold-
I
Heater
I
Mold Open Formed Sheet
I
PI
Mold Closed Fig. 24.24 Schematic of flow molding, flow forming or squeezing flow of nonwoven fiber-reinforced thermoplastic resin composite. (Redrawn from Ref. 95 by permission of Society of Plastics Engineers.)
Other processes forfilled and reinforced polymers 549 Compression molding is also used with halves close on the parison, pinching it long fibers or continuous fibers. In one exam- between the mold halves. The remaining porple, a fiber preform is inserted in the mold tion of the parison is then inflated against the cavity and a heated neat polymer preform is mold walls (Fig. 24.25). Parison thickness conplaced on top. The press is closed to 35 MPa or trol is the key to uniform wall thickness. more, squeezing the polymer into the fiber pre- Parison thickness uniformity is governed by form. The process works best if the polymer is extrudate swell as the polymer exits the die crystalline and has a very low melt viscosity at and by parison sag owing to the parison the molding temperature. Compression hanging weight. The former is a function of molded composite parts are usually more com- the viscoelasticity of the polymer in general plex than thermoformed composite parts. and the normal stress difference in particular. Voids, warping, fiber prominence at the part The latter is a function of the mass of the parisurface intially against the mold and resin rich- son, the extrusion time and the elongational ness at the other surface are typical problems viscosity of the polymer. Neat polyethylenes attributed to the forming process. exhibit extensive extrudate swell. Fillers and fibers reduce the extrudate swell in proportion to the filler or fiber loading (Fig. 24.7 OTHER PROCESSES FOR FILLED AND 24.91°1J02). Filler loading increases parison difREINFORCED POLYMERS ferential weight. And fillers and fibers cause As noted above, foamable, filled and shortfiber reinforced thermoplastics are usually processed in fashions similar to the neat therMold Half Mold Half Parison moplastic. Filled and reinforced hollow structural parts are fabricated by blow molding and rotational molding. Rotational molding is also used to produce multilayer structures having foam cores97. Polyolefins account for approximately 90% (wt) of all non-disposable industrial products produced by blow molding and nearly all products produced by rotational molding. Mica flake reinforced HDPE to 30% (wt) has Mold Open been accumulator blow molded into flat struc380" Pinch-Off tural parts and ducts for automotive, truck and agricultural vehicles since the early 1 9 6 0 ~ ~ ~Graphite-filled -'~~. HDPE is blow molded into conductive electrical boxes and doors. Glass fiber-reinforced HDPE is blow molded into flotation devices, sailboards and kayaks. Accumulator blow molding machines are used to extrude a large quantity of plastic in a short period of time. The filled or shortFormed Part fiber reinforced polymer is plasticated at a constant rate in a conventional screw Mold Closed extruder. The melt is stored in an accumulator until the mold is cleared, then extruded in sec- Fig. 24.25 Top view schematic of lay-flat squeezing onds into a parison or vertical tube. The mold of extruded parison in structural blow molding.
550 Processing thermoplastic composites Table 24.16 Polymer properties important in blow molding Shear and temperature dependent viscosity over the shear rate range of 0-10 000 s-' Melt fracture potential Nonisothermal melt strength or temperature dependent elongational viscosity Extrudate swell Rheological characteristics Normal stress difference Temperature dependent elongational viscosity Polymer strain recovery Molecular weight dependency Molecular weight distribution dependency Particulate polymer frictional coefficients Solid and liquid polymer thermal properties Thermal diffusivity Thermal conductivity Heat capacity Particulate bulk density Pressure and temperature dependent melt density Polymer shear sensitivity Degradation potential - temperature limitation Effect of orientation on gas permeability of polymer Strain-oriented crystallinity levels and effect on Barrier properties Permeability Tensile strength
I
i!
II
II
tlne Shafia
C h d n Drira
Fig. 24.26 Schematic of co-axis rotational molding for filled or lightly reinforced p~lyolefins'~~.
Other processes for filled and reinforced polymers 551 the parison to cool more quickly. As a result, parison wall thickness control must be changed radically when blow molding filled or reinforced polymers. Table 24.16 gives a list of polymer properties important in blow molding. Rotational molding is an atmospheric process in which polymer powder is charged to a metal clam-shell mold. The mold is rotated about the polar or major and equatorial or minor axes (Fig. 24.261°3),while being heated in a forced air convection oven for several minutes, until the powder sticks to the mold surface, melts and densifies into a void-free hollow object. The mold is then air- and/or water mist-cooled, the part removed, the mold recharged and the process repeated. Polyethylene is the major rotational molding polymer, with particle sizes ranging between 50 pn and 500 pm. Coarse particle fillers having particle sizes of about 50 pm or more, such as CaCO,, milled glass and glass cullet are successfully molded to loadings of 30% (wt). Fine particle fillers such as TiO,, carbon black and talc fluidize readily in the tumbling environ-
ment of the mold and so rarely yield useful products. Fibers with aspect ratios of 1000 or less can be successfully molded if the maximum fiber content is less than about 15% (wt). Lower fiber loadings may be necessary for certain mold geometries and certain types of polyethylenes. If the fiber loading is too high, the fibers orient at right angles to the mold surface, producing an unacceptable setaceous inner surface104.One method of forming a hollow composite is to impregnate nonwoven fiber mat with up to 30% (wt) electrostatically charged polymer powder, then sinter the structure to fuse the powder to the fibers. The hot structure is then manually pressed against the mold surfaces prior to adding additional polymer powder, closing the mold, rotating it, heating it to the forming temperature, and cooling it in standard fashion105.Evacuating the mold through the rotating concentric shafts is helpful in minimizing voids but is usually quite difficult to achieve successfully. Table 24.17 gives a list of polymer properties needed for rotational molding.
Table 24.17 Polymer properties important in rotational molding Particle size distribution Temperature dependent thermal properties of particulate polymer Thermal diffusivity Thermal conductivity Heat capacity Bulk density Polymer density Molecular characteristics Molecular weight Molecular weight distribution Crystallization kinetics Heat of crystallization Rate of crystallization Crosslinking characteristics Reaction rate Level Zero-shear viscosity at melt temperature
552 Processing thermoplastic composites
Table 24.18 Effect of fillers on processing properties Processing property
Effect of filler
Viscosity Melt flow Compounding Processing temperature Temperature Injection pressure Flow in injection mold Injection mold shrinkage Injection mold cycle time Thermoforming sheet sag Thermoforming pressure Thermoforming depth of draw Thermoforming part surface quality Extrudate surface quality Extrusion die pressure Melt extrudate swell Melt fracture severity Foam cell size
Increases Decreases Machine dependent, usually complicates the process Increases No effect or cools faster Increases Mold dependent but usually decreases Decreases No effect Decreases Increases Decreases Decreases Decreases Increases Decreases, sometimes dramatically Decreases No effect to decreases
24.8 SUMMARY
Filled polymers tend to process in manners similar to their neat polymer counterparts. The general trend is to increase the melt viscosity of the polymer at low shear rates and to decrease the relative effect of viscoelasticity of the polymer. Table 24.18 shows this for most of the processes described above. Short-fiber reinforcements show more local flow orientation than fillers but, by and large, process parameters are not dramatically influenced by their presence. Long-fiber reinforcements on the other hand show substantial local flow orientation in all polymer melt processes. This orientation is controlled to a limited extent by the design of the dies and molds. Table 24.18 summarizes many of the important processing variables for filled and discontinuous fiberreinforced polymers. Flow of continuous fiber-reinforced polymers is restricted to local squeezing flow around the fiber bundles and so thermoforming, compression molding, stamping and diaphragm forming are the major ways of forming these thermoplastic composites into useful products.
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2.
3.
4.
5. 6.
7.
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10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23.
24. 25. 26.
Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, Table 5.5. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, Chapter 5. Throne, J.L., Plastics Process Engineering, New York: Marcel Dekker, 1979. Middleman, S., Fundamentals of Polymer Processing, New York: McGraw-Hill Book Co., 1977. Tadmor, Z. and Gogos, C.G., Principles of Polymer Processing, New York: WileyInterscience, 1979. Crawford, R.J., Plastics Engineering, 2nd Edn, Oxford: Pergamon Press, 1987. McCrum, N.G., Buckley, C.P. and Bucknall, C.B., Principles of Polymer Engineering, Oxford: Oxford University Press, 1988. Bird, R.B., Armstrong, R.C. and Hassager, O., Dynamics of Polymeric Liquids. Volume 1: Fluid Mechanics, New York: John Wiley & Sons, 1977. Larson, R.G., Constitutive Equations for Polymer Melts and Solutions, Boston: Buttenvorths, 1988. Han, C.D., Rheology in Polymer Processing, New York Academic Press, 1976. Ferry, J.D., Viscoelastic Properties of Polymers, New York: John Wiley & Sons, Inc., 1961. Lenk, R.S., Plastics Rheology: Mechanical Behaviour of Solid and Liquid Polymers, New York Wiley Interscience, 1968. Middleman, S., The Flow of High Polymers: Continuum and Molecular Rheology, New York: Wiley Interscience, 1968. Brydson, J., Flow Properties of Polymer Melts, New York Van Nostrand Reinhold, 1970. Han, C.D., Multiphase Flow in Polymer Processing, Academic Press, 1981, Fig. 3.14, p. 101. Herschel, W.H. and Bulkley, R., Measurement of consistency as applied to rubber-benzene solutions, Proc. Amer. SOC.Test. Mater., 1926,26, 621-674. Hershel, W.H. and Bulkley, R., Kolloid-Z., 1926, 39,291-299. Han, C.D., Multiphase Flow in Polymer Processing, London: Academic Press, 1981, Fig. 3.22, p. 106. Han, C.D., Multiphase Flow in Polymer Processing, London: Academic Press, 1981, Fig. 3.24, p. 108.
27. Shenoy, A.V., Saini, D.R., and Nadkari, V.M, Rheograms of filled polymer melts from meltflow index, Polym. Comp., 1983,453-63. 28. Han, C.D., Multiphase Flow in Polymer Processing, London: Academic Press, 1981, p. 113. 29. Tanaka, H. and White, J.L., Experimental investigations of shear and elongational flow properties of polystyrene melts reinforced with calcium carbonate, titanium dioxide, and carbon black, Polym. Eng. Sci., 1980,20,949-956. 30. Chan, Y., White, J.L. and Oyanagi, Y., A fundamental study of the rheological properties of glass-fiber-reinforced polyethylene and polystyrene melts, J. Rheol., 1978,22,507-524. 31. Kamal, M.R. and Mutel, A.T., The prediction of flow and orientation behavior of short fiber reinforced melts in simple flow systems, Polym. Compos., 1989,10,337-343. 32. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, p. 165. 33. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, p. 166. 34. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, p. 161. 35. Throne, J.L., Thermoplastic Foams, New York Chapman & Hall, 1994, Fig. 9.108. 36. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich Carl Hanser, 1993, Fig. 5.3. 37. Rauwendaal, C., Polymer Extrusion, Munich: Carl Hanser, 1986. 38. White, J.L., Twin Screw Extrusion: Technology and Principles, Munich: Carl Hanser, 1990. 39. Michaeli, W., Extrusion Dies: Technology and Principles, Munich Carl Hanser, 1984. 40. Throne, J.L., Thermoplastic Foams, Hinckley, OH: Shenvood Publishing, 1996, Section 5.2. 41. White, J.L., Twin Screw Extrusion: Technology and Principles, Munich: Carl Hanser, 1990, Chapter 7. 42. White, J.L., Twin Screw Extrusion: Technology and Principles, Munich: Carl Hanser, 1990, p. 10. 43. Rauwendaal, Polymer Extrusion, Munich Carl Hanser, 1986, p. 462. 44. Throne, J.L., Thermoplastic Foams, Hinckley, OH: Shenvood Publishing, 1996, Table 5.4.
554 Processing thermoplastic composites 45. White, J.L., Twin Screw Extrusion: Technology and Principles, Munich: Carl Hanser, 1990, Chapter 11. 46. Rauwendaal, Polymer Extrusion, Munich: Carl Hanser, 1986, p. 403. 47. Wall, D., The processing of fiber reinforced thermoplastics using co-rotating twin screw extruders, Polym. Compos., 1989, 10, 98-102, Fig. 2. 48. Goettler, L.A., Mechanical property enhancement in short-fiber composites through the control of fiber orientation during fabrication, Polym. Compos., 1984, 5,60-71. 49. Michaeli, W., Extrusion Dies: Design and Engineering Computations, Munich: Carl Hanser, 1984,110-115. 50. Han, C.D., Multiphase Flow in Polymer Processing, London: Academic Press, 1981, p. 109, Fig. 3.26. 51. Goettler, L.A., Mechanical property enhancement in short-fiber composites through the control of fiber orientation during fabrication, Polym. Compos., 1984, 5,137-154, Fig. 12. 52. Goettler, L.A., The extrusion and performance of plasticized poly(viny1 chloride) hose reinforced with short cellulose fibers, Polym. Compos., 1983,4,249-255, Fig. 1. 53. Rubin, I.I.,Injection Molding: Theory and Practice, New York Wiley Interscience, 1972, p.2. 54. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, Chapter 5. 55. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, Fig. 5.19. 56. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich Carl Hanser, 1993, Fig. 5.21. 57. Throne, J.L., Plastics Process Engineering, New York: Marcel Dekker, 1979, Chapter 14, 699-800. 58. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich: Carl Hanser, 1993, p. 97. 59. Malloy, R.A., Plastic Part Design for Injection Molding: An Introduction, Munich: Carl Hanser 1994, Fig. 2.77. 60. Malloy, R.A., Plastic Part Design for Injection Molding: An Introduction, Munich: Carl Hanser,
1994, p. 55. 61. Bay, R.S. and Tucker, C.L., Fiber orientation in simple injection moldings. Part I: theory and numerical methods, Polym. Compos., 1992, 13, 317-331. 62. Givler, R.C., Crochet, M.J. and Pipes, R.B., Numerical prediction of fiber orientation in dilute suspensions, J. Compos. Mater., 1983, 17, 330-343. 63. Kenig, S., Fiber orientation development in molding of polymer composites, Polym. Compos., 1986,7,50-55, Fig. 4. 64. Kenig, S., Fiber orientation development in molding of polymer composites, Polym. Compos., 1986, 7,50-55, Fig. 5. 65. Kenig, S., Fiber orientation development in molding of polymer composites, Polym. Compos., 1986,7,50-55, Fig. 1. 66. Darlington, M.A. and Smith, A.C., Some features of the injection molding of short fiber reinforced thermoplastics in center sprue-gated cavities, Polym. Compos., 1987,8,16-21, Fig. 4A. 67. Bay, R.S. and Tucker 111, C.L., Fiber orientation in simple injection moldings. Part 11: experimental results, Polym. Compos., 1992, 13, 332-341. 68. Advani, S.G. and Tucker 111, C.L., The use of tensors to describe and predict fiber orientation in short fiber composites, J. Rheol., 1987, 31, 751-784. 69. Advani, S.G. and Tucker 111, C.L., Closure approximations for three-dimensional structure tensors, J. Rheol., 1990, 34,367-386. 70. Malloy, R.A., Plastic Part Design for Injection Molding: An Introduction, Munich Carl Hanser, 1994, Fig. 2.46. 71. Malloy, R.A., Plastic Part Design for Injection Molding: An Introduction, Munich: Carl Hanser, 1994, p. 55. 72. Malloy, R.A., Plastic Part Design for Injection Molding: An Infroduction, Munich Carl Hanser, 1994, p. 54. 73. Brahmbhatt, S. and Malloy, R., An evaluation of weld line strengths for long glass fiber reinforced styrene/maleic anhydride copolymer, SPE ANTEC Tech. Papers, 1992,38,2563-2567. 74. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich Carl Hanser, 1993. 75. Throne, J.L., Thermoforming, Munich: Carl Hanser, 1987. 76. Throne, J.L., Thermoforming, Munich: Carl Hanser, 1987, pp. 34-35.
References 555 77. Anon., 1900 ultrahigh molecular weight polymer compression molding techniques, Bulletin HPE-102, Himont U.S.A., Inc., Wilmington DE, undated. 78. Narkis, M. and Rosenzweig, N., Eds., Polymer Powder Technology, Chichester: John Wiley & Sons, 1994. 79. Fong, L., Xu, J. and Lee, L.J., Preforming analysis of thermoformable glass fiber mats deformation modes and reinforcement characterization, Polym. Compos., 1995,15,134-146. 80. Throne, J.L., Thermoforming, Munich: Carl Hanser, 1987, Chapter 4. 81. Domininghaus, H., Plastics for Engineers: Materials, Properties, Applications, Munich: Carl Hanser, 1993, Fig. 523, p. 564. 82. Domininghaus, H., Plastics for Engineers: Materials, Properties, Applications, Munich Carl Hanser, 1993, Fig. 457, p. 506. 83. Domininghaus, H., Plastics for Engineers: Materials, Properties, Applications, Munich: Carl Hanser, 1993, Fig. 454, p. 506. 84. Domininghaus, H., Plastics for Engineers: Materials, Properties, Applications, Munich Carl Hanser, 1993, Fig. 457. 85. Bigg, D.M., Hiscock, D.F., Preston, J.R. and Bradbury, E.J., Thermoplastic Matrix Sheet Composites, Polym. Compos., 1988,9,222-228. 86. Wolpert, V.M., Synthetic Polymers and the Paper lndusty , San Francisco: Miller Freeman, 1977. 87. d’A. Clark, J., Pulp Technology and Treatment for Paper, San Francisco: Miller Freeman, 1978. 88. Cakmak, M. and Dutta, A., Instrumented thermoforming of advanced thermoplastic composites. 111: Relative performance of various prepregs in forming double curvature parts, Polym. Compos., 1991,12,354-369, Fig. 3. 89. Throne, J.L., Thermoforming, Munich: Carl Hanser, 1987, p. 28, Fig. 1.22 90. Berins, M.L., SPI Plastics Engineering Handbook of the Society of the Plastics Industry, lnc., 5th Edn, New York Van Nostrand Reinhold, 1991, Figure 13-llb, p. 392. 91. Dutta, A, Niemeyer, M. and Cakmak, M., Thermoforming of advanced thermoplastic composites. I: Single curvature parts, Polym. Compos., 1991,12,257-272. 92. Cakmak, M. and Dutta, A., Instrumented thermoforming of advanced thermoplastic composites. 11: Dynamics of double curvature part formation and structure development from PEEK/carbon fiber prepreg tapes, Polym. Compos., 1991,12,338-353.
93. O’Bradaigh, C.M. and Pipes, R.B., Issues in diaphragm forming of continuous fiber reinforced thermoplastic composites, Polym. Compos., 1991,12,246-256. 94. OBradaigh, C.M. and Pipes, R.B., Issues in diaphragm forming of continuous fiber reinforced thermoplastic composites, Polym. Compos., 1991,12,246-256, Fig. 2a-2b. 95. Bigg, D.M., Hiscock, D.F., Preston, J.R. and Bradbury, E.J., Thermoplastic matrix sheet composites, Polym. Compos., 1988, 9, 222-228, Fig. 6. 96. Progelhof, R.C. and Throne, J.L., Polymer Engineering Principles: Properties, Processes, and Tests for Design, Munich Carl Hanser, 1993, pp. 471475. 97. Shutov, F., lntegral/Structural Polymer Foams, Berlin: Springer Verlag, 1986, Chapter 10. 98. Peters, D.L., Kowalski, R.C. and Hughes, J.K., Blow molded reinforced HDPE for structural applications, SAE Tech. Paper No. 830077, 1983. 99. Rosato, D.V. and Rosato, D.V., Eds., Blow Molding Handbook: Technology, Performance, Markets, Economics. The Complete Blow Molding Operation, Munich: Carl Hanser, 1989, p. 481. 100. Rathgeber, J., Evolution of an imaginative technology - double wall blow molding, in Blow Molding Handbook: Technology, Performance, Markets, Economics. The Complete Blow Molding Operation, Munich Carl Hanser, 1989 (D.V. Rosato and D.V. Rosato, Eds), pp. 828-829. 101. Han, C.D., Multiphase Flow in Polymer Processing, London: Academic Press, 1981, Figure 3.26, p. 109. 102. Minagawa, N. and White, J.L., The influence of titanium dioxide on the rheological and extrusion properties of polymer melts, J. Appl. Polym. Sci., 1976, 20, 501-523. 103. Throne, J.L., Rotational molding, in Polymer Powder Technology, (M. Narkis and N. Rosenzweig, Eds.), Chichester: John Wiley, 1994, Fig. 3. 104. Ramazzoti, D.J., Rotational molding, in PZastzcs Engineering Handbook of the Society of the Plastics Industry, Inc., (J. Frados, Ed.), 4th Edn, New York Van Nostrand Reinhold, 1976, p. 353. 105. Crawford, R.J., Rotational moulding of plastics, Prog. Rubb. Plast. Technol., 1990,6(1),1-29.
TOOLING FOR COMPOSITES
25
Jerry L. Cadden and Paul F. Sadesky
lized for tooling, but no one material solves all The manufacture of composite details and of the problems, particularly when factors such as cost, longevity and tolerances are conassemblies requires that some kind of accurate sidered. The primary objective of any tool for repeatable tool surface be provided, indexed composite fabrication is to make an accurate to an engineering database or reference model repeatable part, within the confines of the and be capable of withstanding repeated exposures to the cure cycle environment of high process parameters defined by the composite temperatures and pressures. Once the specific material supplier and the detail performance manufacturing process has been selected (i.e. characteristics of the end use customer. Design vacuum bag lay-up or resin transfer molding), of the initial tool becomes the most pressing initial issue of tooling for composites. decisions regarding tolerances, heat up rates, coefficients of thermal expansion, tool longevity etc. influence the construction of the 25.2 TOOL DESIGN BASICS tool from an engineering design and material selection standpoint. Individual composite 25.2.1 COEFFICIENT OF THERMAL EXPANSION parts or details will require a variety of support tooling beyond the initial cure tool, such One of the most critical parameters in the as master model reference patterns, trim or design of tooling for composites is the differrouter tools, precision hole location drill tools, ence between the coefficient of thermal assembly fixtures, ply locating templates and expansion (CTE) of the tool being designed other shop aids. Planning must ensure that a and of the composite detail being fabricated. point of reference is established that will con- During the cure cycle of the composite lay-up trol all tooling in any one part family. This will on a tool, the lay-up expands during the heat guarantee that critical dimensional tolerances up cycle. The specific rate of expansion is are maintained within the relationship directly related to the type and combination between different tools supporting the fabrica- of resin or matrix and fibers or reinforcement tion of one composite detail or assembly. In used. The tool will also expand and contract addition, coordination between various com- at a specific rate determined by the material posite details will ensure that and construction techniques utilized. If the interchangeability or replacement is main- CTE values for the laminate and the tool diftained throughout entire structures. There is fer significantly, stresses may result in the an extensive list of materials which can be uti- laminate causing the occurrence of dimensional, strength and part stability problems. The greater the difference between the CTE of the composite detail and the tool, the more Handbook of Composites. Edited by S.T. Peters. Published in 1998by Chapman & Hall, London. ISBN 0 412 54020 7 pronounced the effect will be. 25.1 INTRODUCTION
Tool design basics 557 One of the effects that occurs as a function of these dimensional differences is called springback. Composite details, when cured, hold the specific molded shape, as defined by the tool, as a result of the cured combination of resin and reinforcement. The springback, or more accurately defined as a warpage condition, occurs when the composite detail is cured into a tool, that at a specific temperature has one definite dimensional tolerance and then upon cooling to ambient temperature, contracts to its original ambient dimensions. The composite detail, based on the resin chemistry, cures during the specific period when tool expansion is at it greatest. Warpage occurs when stresses are induced to the composite as the tool begins to return to the ambient dimensions, because the composite detail is being forced to conform to the new dimensional range against the dynamics of the state it reached during cure. This condition will increasingly become greater as the temperature difference between ambient and cure temperature increases and the dimensional size of the tool increases. A common method of minimizing the effects of springback or warpage of quasi-isotropic composite details during and after cure cycling is to determine the CTE of the composite part being fabricated and the CTE of the tooling material selected. During the design of the tooling, carefully match as closely as possible the appropriate tooling material CTE to that of the composite detail. Other conditions that might lead to a warpage of the laminate include an unbalanced laminate orientation where the number of layers or plies of material are more dominant in one direction than another. This condition is separate from any function of the tool and must be considered during the design of composite detail. 25.2.2 USING CTE IN THE DESIGN OF TOOLING FOR COMPOSITE
Two methods are commonly used to minimize the effect of CTE when designing tooling for
the fabrication of composite details. One method is careful selection of the appropriate tooling material. Each of the commonly used tooling materials available has a specific CTE value (Table 25.1). When selecting the appropriate tooling material based on the issue of CTE compatibility only, first determine the CTE of the composite detail being fabricated. The specific expansion rate will be determined by the combination of the resin and reinforcement utilized along with the particular fiber orientation that is incorporated into the laminate. For example, the CTE of the more common unidirectional carbon fibers used in most composite epoxy laminates is approx. 4.5 x 104/OC (2.5 x 10"/OF ). The strength of the reinforcement material lies along the direction of the fiber, not perpendicular to it. If a laminate is balanced, quasi-isotropic, with individual laminate layers or plies equally distributing loads throughout the laminate, the CTE of the laminate will be consistently equal in all directions. If one direction is dominated by more material plies than any other direction, the CTE value will vary, with the dominant direction having the lower CTE. Once the CTE of the laminate is determined, using the appropriate chart select the tooling material with the closest match to the laminate value. The other method of accounting for CTE variations between the detail being fabricated and the associate tool, besides material selection, is the use of shrink factors in the calculation of dimensions prior to tool fabrication. If requirements dictate that when fabricating the tool, a material with an incompatible CTE to the detail being manufactured must be used, steps may be taken to minimize the effect of this variation. During the design phase of the tool, accurate estimates of the actual tool size at its greatest expansion point or at the highest temperature during the cure cycle must be made. The percentage difference between this calculation and the dimensions at ambient must be applied to the base design as a 'shrink factor' reducing the
558 Toolingfor composites Table 25.1 Properties of commonly used tooling materials
Coefficient of thermal expansion, x 10-fipc
Density, g/cm3
(x 1 c 6 / a F )
(iwt3)
Thermal conductivity, cal cm/s cm2"C (BTU in/hft2
Aluminum
23.7 (13.6)
2.7 (170)
0.48 (1400)
Electroformednickel
13.3 (7.4)
8.6 (540)
0.19 (564)
Carbon/epoxy prepreg
4.5 (2.5)
1.61.5 (87-94)
16 x 10-4 (4.6)
Glass/ epoxy
13.1 (74
1.6-2.0 (100-125)
6x (1.7)
Invar 36
3.4 (1.9)
8.0 (504)
0.016 (48)
Invar 42
5.4 (3.0)
8.1 (507)
0.016 (48)
Monolithic graphite
2.16 (1.2)
1.76 (110)
0.28 (840)
High carbon cast steel
9.7 (5.4)
7.3 (456)
0.12 (360)
12-14 6.7-7.8
7.8 (490)
0.10 (300)
22 (12)
0.7
(44)
0.90 2600
50 (28)
0.77 (48)
Material
Steel Wood (mahogany) Urethane board stock
size of the tool by that same percentage. For example, if a specific tool at 177°C (350°F) has a growth factor equaling 1.27mm (0.050 in) of growth over ambient dimensions, this same factor would be applied as a reduction of the overall tool dimension while at room temperature. This method of applying a shrink factor to allow for variations of CTE between tools and details must be approached cautiously when complex shaped surfaces are involved. This is due to the potential for the composite details to become entrapped within the geometry of the tool as the tool returns to the ambient temperature dimen-
O F )
n.a.
sions. Since the composite detail was cured while the tool was at the larger dimension, if the detail is confined to the tool surface or restricted from movement due to complexity in the tool surface, the composite detail could become entrapped, resulting in dimensional abnormality in the laminate and possible damage to the composite detail or the tooling surface. A variety of other factors should also be considered such as tool durability, tool usage rates, thermal conductivity of the material and machinability or fabrication cost. Each of these factors must be weighed individually before final selection is made.
Tool design basics 559 25.2.3 MATERIAL CHOICES IN THE DESIGN OF TOOLING FOR COMPOSITE
The use of CTE is not the sole determinant for selection of a tooling material. For example, as defined in Table 25.1 monolithic graphite which as a tooling material has a CTE value of 2.2 x 104/OC (1.2 x 1O4/'F) has good machinability at a relatively low cost compared to other tooling materials, but exhibits poor durability when utilized in the high usage of production environments. InvaP 36, however, displays a similarly low CTE value, displays a much higher level of damage resistance but has poor machinability and also has a high acquisition cost. Careful selection of the appropriate material for tool use must include review of the following criteria: 0 0 0 0 0
anticipated tool usage (expected life of tool); cost available for tool fabrication; materials available for tool construction; available methods of tool manufacturing; level of dimensional tolerances required from composite detail.
Anticipated tool usage The life expectancy of any tool fabricated for the lay-up and cure of composite details is dependent on a variety of factors. Material selection, shop handling procedures and cure cycle times all affect the ability of the tool to withstand long usage. Certain materials display characteristics that allow longer tool life, however the advantage and disadvantage of each material must be analyzed prior to selection. In addition, each of the tooling materials presently used is sensitive to damage specific to that material. If short term usage is anticipated, temporary tooling such as wet lay-up epoxy or polyester tooling (dependent on detail cure temperature) may be acceptable, however master models and intermediate transfer tooling would be necessary to maintain the correct surface tolerances. To minimize the costs associated with master models and intermediate transfer tooling,
machining the final tool surface directly from a computer model can be accomplished using a material capable of the final cure temperatures. Monolithic graphite and a variety of epoxy and polyester tooling boards would allow this with monolithic graphite offering the highest quality and lowest CTE at a comparable cost to the board stocks available. These costs would be similar to laminated tooling without the intermediate and time consuming steps necessary to complete the laminated tool. However, in a high usage production program, these materials can be damaged more easily than the composite laminate or metallic tools currently in use. Continued advances by the suppliers of composite tooling prepregs have drastically increased the ability of composite laminate tools to support a high number of cycles at cure temperature. Another factor directly affecting the longevity of tooling fabricated from composite materials is the effect of proper employee training in the care required of such tooling. Improper handling techniques will drastically shorten the expected life of composite tooling. Employee caused damage such as cutting on the surface of the tool, using sharp instruments to facilitate the removal of details after cure and improper application of release agents are the most readily identified causes of shop induced damage. Special handling procedures and employee indoctrination can minimize this type of damage. While all tooling, both metallic and non-metallic, is susceptible to damage, tooling fabricated from composite materials is especially sensitive to surface damage caused by employee carelessness. v
Cost available for tool fabrication Cost of tool fabrication is difficult to quantify since material procurement and labor cost vary widely throughout the industry. However, comparisons can be made between the different tooling materials and methods
560 Toolingfor composites up fiberglass or graphite tooling. Equipment obligations can be as minimal as a calibrated surface table and measuring instruments for the accurate setup and weighing of mixed resins and a general knowledge of the systems involved. Use of composite prepreg materials requires additional employee skills levels plus the additional equipment expenditures of an oven or autoclave to cure the materials according to manufacturers’ instructions. Increasingly stringent facility requirements involving the installation of a controlled environment, such as a ’clean room’, have been recommended by material suppliers. As the complexity of the tooling rises, so does the requirement for adequate employee training and engineering support. To support the fabrication of complex, highly accurate machined tooling, machine tools with sufficient work envelopes are required in the correct axis of motion to support the complexity of tooling surfaces. Three-axis machine tools will support basic Materials available for tool construction simple contours, with 4,5 and 6 axis machines Because of the advent of computer-aided capable of machining the much more complex design and manufacturing systems compound contour surfaces. A sufficient com(CAD/CAM) to support tool fabrication, there puter-aided design and manufacturing are more materials now available. Baseline (CAD/CAM) system is required to supply methods such as plaster master models and software commands in order to drive machine plastic-faced plaster transfer tooling will tool directional and cutting speeds. In addialways have a specific application within tool- tion, if board stock materials are chosen for ing, but as the use of CAD/CAM increases reference patterns, the skill and equipment beyond aerospace into all regimens (e.g. sport- necessary to assemble, bond and seal the ing goods, medical and transportation), such material after machining are required. tooling practices that depend more on the skill level of shop personnel than on the accuracy of Level of dimensional tolerances required a machine tool are rapidly being replaced. from composite detail
of manufacturing to assist the user in determining the optimum approach based on a cost/performance comparison. Comparisons must include all necessary steps to deliver a completed tool family. There is a temptation to compare tooling cost for just the lay-up mold in vacuum bag processing and ignore the procedures and tooling necessary to obtain that final surface. In direct comparison, a machined lay-up tool will appear to be substantially more expensive when compared to a lay-up tool fabricated from a composite prepreg. However, when the associated costs of reference patterns, additional splashes to obtain the correct surface and possible autoclave or oven processing time, are factored in, the tool machined directly from numerical control (NC) data, which will eliminate the need for all the intermediate steps mentioned, will become more comparable in price.
Available methods of tool fabrication Methods of manufacturing vary, dependent on equipment and personnel resources available. Plaster type master models, intermediate transfer tooling and wet lay-up type molds take minimum facility requirements. Basic shop skill levels require a familiarity with resins and geometry in order to support simplistic wet lay-
Based on the type of manufacturing method and the type of material selected, different levels of dimensional tolerances are possible. Initially, the designer determines the level of dimensional tolerance for the composite detail being fabricated. Compliance to this tolerance is critical in meeting structural demands and conformance to any form, fit or function requirements.
Tool design basics 561 In the early days of advanced composites, applications were limited to aerospace, which invoked strict dimensional requirements for both tooling and detail parts. These requirements continue within aerospace today. With the expansion of composite usage into other areas, such as sporting goods and automotive applications, the range of acceptable dimensional variations has increased but visual requirements are much more stringent than before. The variation of CTE between different materials for tooling has a major effect on dimensional tolerances. In addition, some materials display sensitivity to environmental conditions that have an adverse effect on dimensional stability. Tooling, such as reference patterns or master models manufactured from urethane board stocks or plaster models, are hygroscopic and may absorb moisture from the atmosphere. This condition will, at a minimum, cause dimensional changes related to the level of moisture absorbed. Also, this condition could be excaberated if the moisture contamination is extensive and the model is taken to elevated temperatures. At higher temperatures, the moisture will expand and may result in possible significant changes in structural integrity. Urethane or epoxy board stock materials have greater resistance to moisture. However, if machined and used for reference patterns they are both still susceptible to contamination, which could result in dimensional changes and possible failures in the bond joints between the block surfaces. Additional steps must be taken to protect all models manufactured from these materials to prevent these types of contamination. Sealers must be applied and the items must be segregated from potential sources of contamination. Monolithic graphite offers advantages over these materials because it is inert and resists contamination from the environment. Lay-up molds manufactured from both ferrous and nonferrous materials must be protected from oxidation. Failure to maintain a nonoxidized lay-up surface will require restoration of the tooling surface which could
result in dimensional changes occurring during this process. The highest level of tolerances available are obtained by machining the tooling surface directly from the computer model. Certain master model materials such as plaster and some board stock materials have a limited temperature exposure level which inhibits the ability to pull composite laminates directly off their surface. Intermediate tooling must be fabricated either to obtain the correct surface level from the master or to be capable of withstanding elevated temperatures within an autoclave or oven, above that of the original model. By machining directly from the engineering database, the need for intermediate surface splashes to obtain the correct surface is eliminated. Each time a splash or the original surface model is duplicated, a 'stackup' or accumulation of the tolerances for each model is combined, resulting in a much greater range of tolerances in the final tool. For example, if a plaster master is fabricated to k0.25 mm (a.010 in) tolerance and each of two additional splashes have the same tolerance range, prior to fabrication of the final lay-up mold, the beginning tolerance range is now d . 7 6 mm (4.030 in). If a tool is machined directly from NC data, the tolerance stackup is eliminated and only the range of the individual machine tool applies. Machine tools, dependent on the condition and environment of machining, are capable of providing 4.128 mm (4.005 in) accuracy or greater. 25.2.4 DESIGNING TOOLS FOR RESISTANCE TO FAILURE
Because of the abusive environment experienced by tooling during the fabrication of composite details, life expectancy of tooling will always be short of anticipation. Repetitive cycling from ambient to over 177°C (350"F), inadequate care and handling procedures, incorrect fabrication techniques have all led to a variety of defects resulting in premature temporary or permanent failure of the tool.
562 Toolingfor composites Failure modes common to composite lay-up tools fabricated by both wet lay-up and prepreg methods generally involve fiber separation. This is due to a variation in CTE between the resin matrix and the fiber. Generally, the neat resin systems used in most composite tooling systems have a CTE of 65 x 104/"C (36 x 104/"F) The graphite fibers used in most prepreg tooling systems have a CTE of around 4.5 x 104/"C (2.5 x 104/"F). During exposure to cure cycles where temperatures will vary from ambient to 177°C (350°F) and above, the difference in CTE between the fibers and the resin will eventually cause disbonds between laminate layers resulting in leaks internally within the tool. In addition, the expansion of the resin is somewhat controlled by the fiber reinforcement in the x and y axis. Because no reinforcement exists in the z plane linking the individual layers together, the difference in CTE between the resin and reinforcement becomes more pronounced. Failures between the individual plies increase because of the lack of reinforcement restraining the resin from the repeated expansion and contraction. Furthermore, when laminating layers of either prepreg or wet lay-up tooling, by cutting each of the plies into pieces 304-457 mm (12-18 in) square, no continuous fiber path will pass through the tool. By discontinuing these pathways, leaks occurring along the fiber path will be minimized. In addition, because each layer or ply consists of sections without any continuous fiber path, stresses within the laminate will be lower, minimizing warpage during use. A majority of failures in composite tooling may be directly tracked to leakage around tooling holes or plumbing fittings. Tooling hole fittings are exposed to repeated shocks during the removal of cured composite details. If steel bushings are used, the difference in CTE will possibly lead to cracks in the tool surface which will become potential leak paths. One solution is to install Invar 36 bushings in laminated tools which are closer in CTE to the parent tool. Also, when
laminating the initial tool, apply additional plies of pregreg in the area of the bushings to increase support in those areas to resist movement during part removal. Another possible solution to the problem of delamination between layers of prepreg tooling is the application of glass transition temperature values (TJ to extend the life expectancy of a tool. Most tooling resin systems are formulated with a Tg value at or slightly above the maximum use temperature of the resin system. As a function of the resin chemistry, glass transition temperatures decay or reduce with each exposure to the cure temperature that the system was designed for. This decay, in incremental steps will continue until well below the cure cycle temperature that the tool was intended to be cycled at. When this point is reached, the resin will begin to break down with a mechanical failure of the bond between the resin matrix and the fiber reinforcement. The solution to this problem is to use a resin system with the highest possible Tg value available. For example, if the tool is intended to be cycled repeatedly at 177°C (350"F), a T value of the resin system in the 220°C (425"h) range will allow more cycles. It is common among some aerospace companies to now fabricate composite tooling for epoxy laminates from a bismaileimide or a cyanate ester resin system with Tg values higher than 260°C (500°F). This allows the inevitable decay of the Tg value to span a greater difference allowing the life of the tool to be extended. The same principle may be applied for any prepreg system from polyesters to the higher temperature resin systems. One consideration in using this method requires the selection of a master model material capable of exposure to the elevated temperatures that the higher temperature systems require during cure. Plaster and most of the board stocks available are not capable of these higher temperatures and intermediate splashes or surfaces would have to be provided. Monolithic graphite does provide a surface capable of exposure to higher
Master models 563 temperatures in addition to having the lowest CTE available, allowing tools fabricated from these resin systems to be taken directly off the model surface without need for intermediate surfaces. Employee-induced damage of tooling can play a much greater part in the reduction of expected tool life. Correct indoctrination into the importance of the tool to the fabrication of accurate details must be stressed. Most of the employee-induced damage will occur either during the lay-up procedures, the removal of cured composite details from the tool surface, or preparation of the tool surface prior to the next lay-up. During lay-up of details, damage will be the result of employees using knives or other sharp objects during the trimming of the composite material. If proper care is not taken, the employee will not only cut the material but also cut into the tooling surface. While not as detrimental to a metallic tool, damage of this kind may be catastrophic to a tool fabricated from a composite material. The cut will allow a breach in the vacuum integrity in addition to allowing resin to penetrate beyond the surface of the tool. Also, when laminates are removed from tooling after completion of the cure cycle, damage occurs when personnel use sharp equipment to force the completed detail from the tool. The greatest care must be given when attempting to remove the detail, to prevent inadvertent damage if the detail fails to cleanly release from the tool surface. Damage, not only to the tool, but also to the detail may result. To prevent this damage from occurring, proper steps must be taken. Employees must be indoctrinated in the proper techniques of tool maintenance and lay-up procedures and must be provided with acceptable tooling aids to assist in the safe removal of cured details from the tool surface. Soft wood or plastic wedges must replace hammers and hard-faced chisels for detail removal and tool surface preparation. Tools must be designed with adequate laminate thickness to prevent damage to the tool if
struck with a hard object. Carriers designed to transport the tool to and from work station must also function as a protective barrier to prevent the tool from striking walls or beams within the shop environment. Support tooling, where applicable, must be designed to be as lightweight as possible to prevent injury to the employee and damage to the tool surface when handled. 25.3 MASTER MODELS
A master model is considered to be just that a master source identified with holes, scribe lines, trim lines or any other feature of the part that requires duplicating to other tools. The master model is the physical representation of the design or a point of reference to which all support tooling, both for fabrication and inspection, would be indexed. Because this surface will provide the reference pattern for all subsequent operations beyond initial fabrication, such as assembly fixtures indexing a variety of details from different locations, extreme care must always be taken in protection of the master model. Master models may be fabricated from a variety of materials. Common materials include plaster, machined urethane or epoxy board stock, monolithic graphite or most ferrous and nonferrous metals. Each material offers distinct advantages and disadvantages. To determine which material is the most feasible, the entire tooling family philosophy must be reviewed. Master models are generally stored indefinitely so that they may be referred to over the life cycle of the manufactured parts. In cases where cost and/or time schedules are important, temporary models are produced and then destroyed once they have been used. However, because of the hygroscopic nature of plaster, care must always be taken to protect the master model from the environment to maintain accuracy. Adequate storage containers, allowing for complete protection, must be utilized throughout the life expectancy of the model.
564 Toolingfor composites 25.3.1 PLASTER MASTERS
25.3.2 TEMPLATE! METHOD
One of the oldest methods of producing a mas- There are several techniques of building a ter is from plaster. Plaster is made from the plaster master determined by the shape of the mineral gypsum (CaSO,). which is finely part. If the part is not symmetrical and does ground and calcined (dehydrated) to produce not have a constant cross section or the size is a fine powder with uniform properties. With large, the master model is made from a series the addition of water to form a workable of templates secured to a flat base to form a slurry, a reaction occurs which produces heat three dimensional full scale model of the part. and the inert gypsum on drying. Plaster is Space between the templates is relative to the manufactured in various textures or grades degree of abruptness of the contour. For norwhich support the level of detail required on mal gentle contours a space of 15.24-20.32nun the model. Coarse grades are used to build up (6-8 in) is common. Templates are usually the master model surface and then followed made from 0.317 mm (0.125 in) thick aluby the fine grades which allow precise details minum to prevent corrosion. For temporary such as trim lines or other identifications to be masters, steel is sometimes used, but, because scribed into the surface. Depending on the of the amount of moisture used in the mixing grade being used, plaster has a setting expan- and application of the plaster, steel templates sion of approximately 0.080% and a thermal may rust (Fig. 25.1). expansion in the dried state of a maximum of If electronic data is available, the templates 0.027/ "C (0.0156/ OF). can be NC machined or cut with a water or
REFERENCE LINE REFERENCE LINE
Fig. 25.1 Skeletal structure for template plaster master.
Master models 565 laser jet directly from the flat pattern generated by the data. Section cuts taken at specific station lines from two-dimensionalblueprints can be used to saw out a template. Except for the NC machining method, deburring is generally required to remove spurs or sharp edges from the templates prior to use. Holes are drilled into the templates for threaded rod spacers and screen support rods. For larger models, air passages are cut into the bottom of the template to allow for even curing of the plaster. Once sufficient templates have been prepared, bluing is applied to a flat ridged steel table and scribed with an awl to denote the location of each template. Flatness of the table is critical and should be within 0.127 mm (0.005 in). Tooling balls which indicate the x, y and z direction are sometimes placed on the table corners as reference points for the system. Tooling balls can vary in size but a common size is 12.7mm (0.5 in) diameter on a 6.35 mm (0.250 in) diameter x 12.7mm (0.5in) long pin. The pins, each with a 'ball' on top are placed into location holes and optically sighted relative to the position of the each ball location. Location can also be treated relative to a position on the master such as a station line. Each template is attached 90" to the base table with angles and held to within 0.127 mm (0.005 in) of the reference line at the base, the face square to the base to within 0.076 mm (0.003 in) in 304.8 mm (12 in) and within 0.127 mm (0.005in) of the base reference line. Threaded rods are secured with sheet nuts on each side of the template to provide rigidity to the template face. Wire mesh is placed between the templates and secured to the threaded rod with wire hooks approximately 101.6 mm (4 in) below the top surface of the template. This is used to hold the plaster in place. Plaster is mixed with hemp and placed against the screen to approximately 9.5 mm (0.375 in) below the template surface.A second layer without hemp is added to this surface to approximately 12.7 mm (0.5 in) A sawtooth scraper is used to build a striated surface and allowed to dry. A final mix is made with the fine grade of plaster and using a flat spring
steel blade, the plaster is 'faired' or swept flush between the templates to form a smooth accurate surface. Because of the propensity of plaster to absorb moisture, it should be sealed after the surfacehas had adequate time to cure. Commercially available lacquers can be used to seal the surface and provide a suitable protection within the shop environment. 25.3.3 FOLLOW BOARD METHOD
A method widely used when a constant cross section is to be built is the follow board. A flat surfaceis required with an accurate side surface to act as a guide rail. A template of the contour is prepared from a rigid 3 mm (0.125 in) minimum sheet of aluminum or steel and attached to a wooden guide support. Plaster is mixed and built up on the surface to within 3mm (0.125 in) of the final contour. Partial drying is recommended before the final plaster mix is applied. This will prevent shrinking and cracking of the plaster surface which would affect accuracy. Using the template and guide support, the plaster contour is formed by pushing the template evenly over the surface (Fig. 25.2).
-
/
FOUOW BOARD
FOLLOW BOARD METHOD FOR PRODUCING PLASTER MASTER
Fig. 25.2 Follow board method for producing plaster master.
25.3.4 SWEEP METHOD
A third method called a sweep is best utilized when a symmetrical surface s;ch as a cone or
566 Toolingfor composites hemispherical shape is involved. As with the follow board method, a flat surface from which a frame can be constructed of the shape to be produced is required. For large shapes, intermittent templates should be placed within the framework to allow support for the sweep and provide adequate support for the sweep to fair against. The sweep itself is usually made from sheet metal 3 mm (0.125 in) minimum thickness and supported by a wooden guide or other mechanical guides that can ride the surface of the flat surface table. Smaller shapes, of course, do not require this extent of rigging. Plaster can be reinforced with saturated hemp fibers, mixed into the slurry and applied to form rough shapes and to form strengthening ribs on the back surfaces of casts. All master models fabricated from plaster require, in addition to sealing with commercial grade lacquer, suitable storage containers if the model is required to be stored for any period of time outside the shop environment.
operation significantly. The system, known as Automated Tool Manufacture for Composite Structures (ATMCS),is an expert system with macros which dramatically speeds up the tool design process. ATMCS takes the composite detail surface model into either IBM Catia or EDS Unigraphics I1 and creates the tool required around the part model. The system, acting through a series of inquiries made to the tool designer, selects the optimal configuration, material, manufacturing process and design. The design is then created around the part model, with significant savings in time. Although the system was developed initially for the aircraft industry and is presently used for basic open-faced layup molds, it could be expanded for many different types of tools and processes such as resin transfer molding and injection molding in other industries. 25.4 COMPOSITE TOOLS
Composite tools are usually made from epoxy resin matrix and either E-glass or carbon fibers Because of the widespread use of CAD (com- as reinforcement. Depending on the life cycle puter aided design) systems, older methods required, tools can be made from prepreg or which utilized two-dimensional prints to by 'wet' lay-up procedures. Prepregs generally build master models are now used less fre- require curing within an autoclave because of quently. With CAD systems, a great deal of the elevated pressure specified by the manuaccuracy can be transferred into the master facturer. Because of the increased compaction model via the NC machining operation. Table available when curing in an autoclave, tooling 25.1 lists various materials widely used today fabricated from prepregs are capable of a for NC machined master models. From a CAD greater number of cure cycles than the wet laymodel of the part, a tool manufacturer must up method. In addition to greater compaction, design a tool from the surface data supplied. autoclave curing offers better control of resin Advances with CAD/CAM systems seek to content and uniformity of reinforcement. minimize the operator input to the system and However, for shop aids such as trim tools, transfer design responsibility to the computer. room temperature curing epoxy systems are One example of this technology is demon- recommended. strated in a system developed by a multicompany team lead by the Northrop 25.4.1 LAY-UP MOLDS Grumman Corporation for the US Air Force Manufacturing Technology Directorate at Lay-up molds are used to form the shape of Wright Patterson Air Force Base. This system, the part to be produced and have the part while not totally removing the tool designer periphery scribed on the surface as well as from the design process, does streamline the the location of any required cross hairs and 25.3.5 NC MACHINING
Composite tools 567 tooling holes. Tools can be made directly from a NC machined master model or from a plastic faced plaster splash taken from a master model not capable of elevated temperatures and pressures. The choice of glass or carbon fiber/epoxy for the mold is generally governed by the complexity and CTE of the part to be fabricated. Lay-up molds must be capable of maintaining a vacuum tight environment while being subjected to high temperatures and pressures. 25.4.2 PREPREG METHOD
There are a considerable number of prepregs available as epoxy 'B' staged glass or carbon reinforced cloth. Prepregs can be obtained in rolls or as precut squares or rectangles. The weave style can vary depending on the amount of drape to be encountered but generally plain or satin weaves are readily available. The resins are tailored for tack, out time and glass transition temperatures at a minimum and are around 40% by volume of the prepreg. (Tooling prepreg manufacturers have very detailed procedures that they recommend for their specific system. These comments are not meant to supersede the recommendation of a manufacturer, but rather to place emphasis on important steps that should not be overlooked for tool fabrication.) Within the last several years, an innovation for tooling prepregs is the capability for low temperature curing 61°C (145°F) in an autoclave, followed by a free standing post cure at 177°C (350°F) after removal of the tool from the master. This has allowed the use of plastic faced plaster and urethane based tooling boards for direct lay-up of composite tools. Monolithic graphite with a low CTE and capability to withstand 315°C (600°F) under autoclave pressures can be a good choice. The first step prior to prepreg application on the master surface is to ensure that the prepreg and the master surface are absolutely clean and free of debris and that the surface is smooth and without pin holes. A quick vacuum check
is always a good idea at a minimum of 6.2 kPa (25 in Hg). A loss of 500 Pa (2 in Hg) within 5 min with the pump nonoperating is acceptable. Apply masking tape around the tool periphery for later application of the sealant tape. It is absolutely necessary, regardless of prior history of the master surface, that it can be released with a suitable hard wax or other release agent. The prepreg manufacturer may recommend a specific release agent for his prepreg system and it is advisable to follow those instructions due to the possibilities of chemical reaction occurring between the resin system of the tool and the release agent used (Table 25.2). After the cleaning and releasing processes have been completed, release coated tooling pins should be placed into the holes of the master. These are generally index and locating holes that have bushings and are used to position or align one tool to another, or to a production part. Bushings can be installed during lay-up of the tool or potted in after final cure of the tool. In general, a face or gel coat layer is not used by most manufacturers today. The reason is associated with the difference in CTE of a neat resin on the face and the CTE of the underlying reinforced prepreg which over the life span of a tool can cause cracking and crazing of the face and subsequent loss of vacuum integrity. If a gel coat is used it should be of minimum thickness to minimize these effects over time. A lightweight (style 7781) cloth is the first layer applied to prevent mark through to the surface from heavier cloth. Carefully lay each ply onto the surface and work out wrinkles or air bubbles and maintain the warp direction of each ply in the 0" direction. An overlap between the plies should be 3-6 mm (0.125-0.250 in). Some manufacturers recommend a debulking step at this point to ensure no air entrapment at the interface and a smooth surface on the tool. Debulking is accomplished by application of a peel ply net to the edge of the laminate and working out wrinkles and air bubbles. A resin dam (sealant
568 Toolingfor composites Table 25.2 Cures for common mold release problems Problem
Cause
Solution
Non-adherance of paint or adhesive to part
Build up of mold release, layers of incompatible mold releases, or too much release in formulation
Use manufacturer’s recommendations for layers and cure schedules. Avoid adding incompatible layers such as silicone, wax and flurocarbon. Check with manufacturer for possible revision or custom formulation to allow multiple releases and adequate paint adhesion
Poor release with small particles of visible on mold surface
Porosity in mold
Thoroughly clean mold with an appropriate solvent and then add a mold sealer before the release coat
No paint or adhesive sticks to part
Multiple release not possible
Check for the presence of silicones Remove silicones from plant in the area. Silicone mold releases where painting or adhesive
can be transferred over long distances
bonding is performed
Poor release in some areas, particularly in severe contours
Apply one or two additional coats of release to severe contour areas
Failure to obtain any release even Inadequate cleaning of the mold before application of release has though release was applied interfered with the ability of the release to bond to the mold or the release has been improperly cured. Shelf life of release may have expired Generally poor release with patches of white transferring to the part
Strip out the part and thoroughly clean the mold, then apply and fully cure release agent. Also check shelf life of release
Mold surface not properly cleaned Remove all release and which results in poor bonding of thoroughly clean mold with an appropriate solvent before release to mold reapplication of release. Follow manufacturer’s recommendations for cure cycle
tape) can be placed around the perimeter to prevent resin flow (Fig. 25.3). Next, lay-up one ply of Teflon@beyond the resin d a m and attach to the resin dam. Using the manufacturer’s recommendation, holes should be placed into the Teflon ply to allow for resin bleed. Pre-perforated film can be obtained for this purpose and provides greater control over the size a n d spacing of the holes. If only one lightweight ply has been applied, no holes are required to
permit a higher resin content on the tool surface. Over this layer, one ply of polyester breather cloth or 7500 style glass cloth is applied. A nylon vacuum bag is placed over the entire stack and a vacuum of at least 625mm (25 inHg) is applied for at least one hour. Removal of the bag, breather, separator film and peel ply should be done very carefully to avoid lift up or shifting of the prepreg layers. The orientation for each additional ply
Composite tools 569
Vacuum Line
Vacuum
Tdl
Laminate
Vacuum Bag Sealant ,,Breather
\
Release Film
Fig. 25.3 Laminate pre-bleed stackup.
should be such that a balanced system is maintained to minimize stress build up in the laminate. Prepreg manufacturers will clearly stipulate the lay-up sequence.After the second or third ply has been laid down, knurled bushings should be placed onto the tooling pins and pressed down to seat them. Subsequent plies will be placed over the bushings to integrated them into the laminate. After the seventh or eighth ply, the pins can be removed so that later plies can cover the bushing completely to prevent vacuum leaks. In some instances a pad or build up of plies over the bushing is recommended. An alternate method is to pot the bushings into the laminate after the final cure. To do this a tapered wax or rubber plug should be placed over the pin to allow space for the potting compound after the final cure. Each ply should be carefully worked into corners and radii making sure all entrapped air is removed. Wrinkles should also be carefully worked out before another ply is placed over it. If a persistent wrinkle or air bubble can not be rolled out, then carefully slit the pockets with a sharp knife and work it down into the surface. In the excess area of the tool, thermocouples can be strategically located between the plies for recording during the autoclave run. As a rule of thumb, debulking should be performed
after every 4-5 plies. Final build up of the laminate should be at least 0.013 mm (0.375 in) or whatever is recommended by the prepreg manufacturer. Final vacuum bagging is performed in the same manner as for debulking with a layer of peel ply, perforated Teflon, polyester breather and vacuum bag. Prepreg manufacturer will provide a detailed heat up rate and cure temperature for their system and this should be carefully followed. Most systems can be initially cured at up to 63°C (145°F) and 586-689 kPa (85-100 psi) of pressure for 14 h. After the autoclave cycle, carefully remove the bag and films from the laminate to avoid lift up from the master surface. Attachment of the egg crate structure (support or back-up structure, Fig. 25.4) to the laminate is very important to minimize any potential residual stresses built into the laminate or stresses from the egg crate itself. Leave the laminate on the master surface and attach board structure of the same material to the surface of the laminate. If the laminate is glass/epoxy, the egg crate material can be made from glass/epoxy or aluminum honeycomb sandwiched between glass. The point is to avoid stresses caused by the difference in CTE between the egg crate and laminate by using similar materials.
570 Toolingfor composites
Fig. 25.4 Support structure details.
The egg crate should be cut to the contour of the laminate with a standoff of 3.17mm (0.125 in). This prevents hot spots during production part curing and also mark off into the tool laminate. The egg crate should be constructed so that it will lie flat on the surface. The intersections of the board stock are held together with cloth and resin with at least three strips per junction. To ensure minimal stress to the laminate from the egg crate structure, it is advisable to remove the structure from the laminate and cure the strips holding the structure together at 177°C (350°F). After curing, the structure can be placed back onto the laminate and 'tied' into the laminate with at least three strips of cloth and resin around the periphery of the egg crate. Shims can be used to provide for the standoff. If the type of master used for the lay-up permits an
oven cure at 177°C (350°F) (Le. metal or monolithic graphite) the post cure can be performed without removal of the tool from the master. However, if the master material will not tolerate this temperature, careful removal of the tool from the master must be done prior to the post cure. Separation of the tool from the master should be done carefully to avoid damage to the master or the tool itself. Tooling pins should be removed prior to separation. Tools with severe contours may require plastic wedges to be inserted around the tool periphery until it releases. Once the tool is separated, the surface should be inspected for pinholes or roughness. Pinholes can be filled with resin and the roughness can be smoothed out with a fine grit sandpaper. Edges of the tool can be sawed to even up the periphery and then
Composite tools 571 sanded lightly with 220 grit sandpaper to occur in corners or the bottom of contours. Too remove any loose fibers. Care must be taken much resin will result in cracking and crazing that no fibers are lifted by sanding along the later in the tool life cycle. Also, ensure that all length of the fibers. air bubbles have been worked out by repeatOnce the tool has been cleaned up, the edly applying the brush back and forth across required check for vacuum integrity is accom- the surface. To ensure that all air has come to plished by placing a layer of polyester the surface while brushing, pause occasionally breather cloth on the surface and a vacuum and allow the air to rise to the surface where it bag over it. The acceptance criterion is gener- can be brushed out. Air that remains ally that there be no loss greater than 500 Pa entrapped either on the tool surface or within (2inHg) in 5 min at a minimum of 6.2 kPa the layers of cloth could result in possible blis(25inHg) at the start of the test. If possible, ters and delamination later during tool usage. depending on the complexity of the tool, place The resin supplier will provide mix ratios for the tool back onto the master and check for resin and hardener as well as pot life and gel any warp or out of contour problems. times. Tooling cloth generally comes in rolls and is either a satin or plain weave with the warp direction noted. Sufficient squares or rec25.4.3 WET LAY-UP METHOD tangles can be precut from the roll prior to The wet lay-up of composite tools can be for laminating. Sections over 609 mm (24 in) room or elevated temperature use. The differ- become too cumbersome to work on the tool ence is in the resin selection. Procedurally, the face, therefore smaller sections are advised. process is the same except for the cure cycles. The first 2-3 plies should be from light weight The master or tooling aid should be cleaned of cloth such as 7500 glass or 2534 carbon which all defects and debris such as scratches and will prevent mark through to the surface. The loose fibers. Solvent clean the surface to gel coat should be advanced with time prior to remove any residual resin or oil. Check for application of the first ply. If enough tack is vacuum integrity using a criteria of a maxi- not present, the ply will sink too deep into the mum loss of 500 Pa (2 in Hg) at a minimum of gel coat and be visible on the tool surface. One simple test is to place a finger onto the surface 6.2 kPa (25 in Hg) starting vacuum. Place masking tape around the periphery and release. If the gel coat has not advanced for later use for sealant tape. Regardless of the adequately, the fingerprint will disappear. If prior history of the tool, it should be released the fingerprint remains, the gel coat has with a suitable hard wax or release agent. advanced far enough to withstand the Tooling pins should be released and placed repeated pressing of subsequent layers of cloth. Using the mix ratios provided by the into the holes provided on the tooling aid. For wet lay-ups, two resins are used, one for supplier, mix enough resin to cover the surface the gel or face coat and one for laminating. The of the gel coat in the time allotted by the pot gel coat is generally the same as the laminating life or around 30min. Approximately 40g resin but with additives to thicken it to make it (0.088 lb) per 0.009 m2 (1ft2)of tool surface for adhere to the contour of the master or tooling each ply should be adequate. Application of aid. Resin manufactures can supply both room the first several plies should be done carefully to avoid pushing through the gel coat surface. temperature and high temperature systems. To ensure complete wetting of the ply, a Apply the gel coat to the surface using a short bristled brush or squeegee. Work the short bristled brush or squeegee can be used to coating as evenly as possible over the surface carefully work the ply into the laminating resin. at a thickness of approximately 12 mm All wrinkles and air entrapped areas should be (0.030 in). Do not allow excessive build up to worked out before another ply is added and if
572 Toolingfor composites necessary, use a sharp knife to slice through the ply wrinkle in order to work it down. Bushings should be placed over the tooling pins at this point and worked into the resin to seat them. A heavier cloth such as 7587 glass or 2548 carbon can be used for the remaining plies. Since the cloth was precut into squares or rectangles and the warp direction was maintained, each ply should be placed at 45" to the previous ply. Overlaps of 6.35 mm (0.250 in) between plies should be maintained but a seam should never be placed over a seam from a previous ply. After each ply, add additional resin to cover the surface. Place the next ply and work the resin up through it by pressing the cloth with the brush. If there is not sufficient resin to completely wet out or saturate the ply, additional resin should be used. Saturation of the cloth with resin on a table separate from the lay-up surface and then transferring the saturated ply to the tool can cause air entrapment and bridging of the ply. After the fourth ply, or prior to the resin curing, apply a peel ply to the surface for a compaction cycle. The peel ply when removed before other operations will eliminate the need for sanding the surface prior to bonding. This is followed by a Teflon film with perforations every 250-300 mm (10-12 in). Cover this with a heavy glass or bleeder cloth to bleed off excess resin during compaction. Finally, place a nylon vacuum bag over the surface using sealant tape to attach to the surface and apply a vacuum of at least 6.2 kPa (25inHg). Hold this vacuum for 10-12h or overnight or until the peel ply can be removed without disturbing the laminate layers. Following this cycle, the bag and peel ply can be removed along with the tooling pins and lamination can commence as previously described. Debulking should be performed after every 6 plies or before the resin begins to cure. Final laminate thickness should be 9.5 mm (0.375in). It is probably a good idea to build up the bushed hole area with additional plies to ensure vacuum integrity. After the final ply has been applied, the compaction step is repeated with the peel ply,
FEP, bleeder and the vacuum bag. Depending on the resin system and the tooling aid material used, a precure is recommended and should be supplied by the manufacturer. Fabricate an egg crate structure using 9.25 mm (0.375 in) thick board stock of similar material to the laminate to avoid stresses caused by the difference in CTE between the egg crate and laminate. If the laminate is glass/epoxy, the egg crate material can be made from glass /epoxy or aluminum honeycomb sandwiched between a glass laminate. The egg crate should be cut to the contour of the laminate with a standoff of 3.17 mm (0.125 in). The standoff prevents heat differences or hot spots on the tool surface during production part curing and also prevents mark off from the back up structure pressing upward into the tool surface laminate. The egg crate should be constructed so that it will lie flat on a surface. The intersections of the board stock are held together with cloth and resin with at least three strips per junction. To ensure minimal stress to the laminate from the egg crate structure, it is advisable to remove the structure from the laminate and cure the strips holding the structure together at 177°C (350°F). This will allow the tool surface to be tied into a stabilized support structure and minimize warpage during subsequent cure cycles. After the support structure is cured, the structure can be placed back onto the laminate and attached or 'tied' to the tool laminate with at least three strips of cloth and resin around the periphery of the egg crate. Shims can be used to provide for the standoff to prevent warpage. If the type of master used for the layup permits an oven cure at 177°C (350°F), then the post cure can be performed without removal of the tool from the master. However, if the master material will not tolerate this temperature, careful removal of the tool from the master must be made prior to the post cure. Allow the tool to stand at ambient temperature for a minimum of 24 h prior to post cure. After the final 177°C (350°F) post cure, inspect the surface for pinholes and repair any
Composite tools 573 blemishes with gel coat resin. A final vacuum check at 635 mm Hg (25 in Hg) with a loss of no more than 51 mm (2 in) is acceptable.
the same thickness and place any bushings through the second layer and flush to the first layer. Place a ply of 7500 glass cloth onto this layer and work in to impregnate the cloth. Allow to cure to the fingerprint test. Mix a 25.4.4 PLASTIC FACED PLASTER third batch of the resin but add about 10-15% Plastic faced plasters (PFPs) are tooling aids by weight of wet plaster to the mix and apply that minimize the wear and tear on masters by to a thickness of 2.5 mm (0.1 in). Do not wait duplicating the master surface with a suitable for curing but proceed with a layer of plaster unit that can be used for a variety of purposes. approx. 25.4 mm (1in) thick. Allow this to PFPs allow for tooling to be directly fabricated partially dry and then finish the tool by from the master surface without exposing the adding plaster and hemp to the surface to a master model to adverse environmental con- thickness that will allow support for the size ditions, such as autoclave temperatures or of the tool [50-76 mm (2-3 in) for a 914 mm x pressures. If taken directly from a master sur- 914 mm (4 ft x 4 ft) tool]. Support structure face, the PFP is the reverse of the master can be built in for small tools using plaster contour. An intermediate plaster splash is and hemp to make strengthening ribs on the required to get back to the master contour back surface. For large tools, steel pipe or tubwith a PFP. If the surface required is directly ing can be tied into the back structure with from the master model, the PFP will be taken plaster and hemp ropes. Approximately 24 h directly from the master surface. If the surface is required to dry and cure the system and is above or below the master surface, appro- depends on the thickness and size. Drying in priate steps must be taken by either taking an oven up to 60°C (140°F) will provide a staadditional splashes with or without layers of ble system for use. PFPs can be used in an tooling wax to achieve the appropriate dimen- autoclave (with vacuum integrity) up to sion. approximately 105°C (220"F), however a limit Prepare the master surface by cleaning thor- of one or two runs is all that can be expected. oughly and removing blemishes, debris and PFPs provide tooling aids for a variety of pinholes by filling with a compatible filler other room temperature shop applications. compound. Release the surface and any tooling pins with a hard wax or release agent. If 25.4.5 DRILL TEMPLATES the master model is plaster, a hard wax can be used with a minimum of three applications, Drill templates or fixtures are used primarily dried adequately and buffed between each to drill and locate precision holes in the proapplication per the manufacturers directions. duction composite part. While their use is From the resin supplier, request a water- limited to hole location and drilling, their proof or hydrophobic resin system which will function may also be combined with other cure in the presence of water from the plaster. support tooling, such as a trim/router fixture Apply the resin evenly 0.76 mm (0.030 in) to minimize tooling expenditures. Drill fixthick to the master surface with a short bris- tures are fabricated using a room temperature tled brush and work out air bubbles as they cured fiberglass/epoxy laminating and face appear. Ensure that no bristles are pulled coat system. Because the tool is used in the from the brush to contaminate the resin shop environment in ambient conditions, no Allow the resin to cure to a point that a fin- vacuum integrity or elevated temperature gerprint may be imprinted lightly into the requirements are needed. Location of the holes resin and will remain for a period of time can be obtained from the master model surafter touching. Then apply a second coat of face. In addition, to facilitate concurrent tool
574 Toolingfor composites manufacturing, a Mylar@film sheet (0.010 in) or thicker may be used as a transfer medium by relocating the position of the holes to the surface of the Mylar and then using the Mylar as a temporary master surface while fabricating the fixture. Prior to fabrication, it must be determined if the fixture is to mount on the outside surface of the part (OML) or the internal surface of the part (IML). Since master models normally represent the OML surface, most support tooling such as drill and trim fixtures may be fabricated directly off this surface. Occasionally the surface required will be a specific distance above or below the surface of the master model. If the surface is above, tooling sheet wax at the specific dimension required must be placed on the master surface prior to fabrication of the tooling aid. Sheet wax is commercially available in numerous thickness to accommodate most requirements. If the surface required lies inside the master model surface, a 'splash' consisting of plaster and hemp reinforcement must first be pulled from the master surface and then the splash surface can be waxed to the specific dimension inside the master surface. Tool pins are placed in the tooling aid. After the face coat and two layers of glass have been applied, the drill bushing is seated onto the surface with additional lamination over the bushing to provide an integral lock to the tool. Template thickness can vary depending on use, but 9.5 mm (0.375 in) in thickness is typical. 25.4.6 TRTM AND ROUTER TEMPLATES
These shop aids are used to trim and rout cured composite parts to a specific dimensional tolerance. Accuracy is required for these tools in order for the composite detail to fit precisely with adjacent details. Trim and router templates can be fabricated directly from the master model, composite tool or a tooling aid such as a PFP (plastic faced plaster). They are generally fabricated using room temperature cured epoxy/glass systems. Since
trim and routing operations are always conducted at ambient temperatures, CTE is not considered in the design of this type of tooling. Procedures for laminating the room temperature cured system are similar to those for drill fixtures. Periodic debulking is not required and the tool is not required to maintain any vacuum requirements. Thickness can vary depending on final use but is usually about 9.5 mm (0.375 in). If the tool is to be used for routing, a set back or offset will have to be determined as defined by the type of routing equipment used. This set back must be identified on the surface of the tool to alert personnel to which equipment is acceptable for use with the tool. Failure to use the correct set back will result in an under trimmed or over trimmed condition. A witness or verification line is usually scribed on trim fixtures as a reference to which edges may be checked for damage. This allows shop personnel to quickly verify the accuracy of the trim fixture with minimal inspection equipment. With drill fixture tooling, a determination must be made as to what surface the trim fixture is applied. If the tool is to represent the same surface of the master model, then the tool may be directly taken from that surface. However, if the surface required is internal or external to the master, appropriate steps including plaster splashes or waxing must be completed to obtain the correct surface. 25.4.7 PLY LOCATING TEMPLATES
Ply locating templates are used during the layup of the production part and designate locations for the plies and indexing of detail parts. In addition, these templates may also show individual ply orientation and designate special features of the part such as splice areas or hardware attachment points. Occasionally, honeycomb core placed within composite production details must be potted with a syntactic core material to prevent core collapse when hardware is attached. Reference locations of all attachment hardware may be transferred
Bibliography 575 from the master model to produce a core potting template. Core may then be accurately potted, laminated and then drilled after curing by referencing the location from the potting template. Similar templates may be produced to locate individual details or other assemblies that have to be located during detail construction. Slits or eyebrows are cut into the laminate to locate the edge of the production part ply and color coded and identified accordingly. All templates are fabricated from room temperature glass/epoxy cured systems and are designed to be light in weight with a thickness of 3.1-3.8 mm (0.125-0.150 in). Because some templates may be quite large, provisions must be made where possible to
lighten the template to assist in handling. Lightening holes can be placed by removal of sections of the template not serving a specific function. However, as material is removed to reduce weight, stiffeners must be added to prevent warpage that may affect dimensional stability. BIBLIOGRAPHY
Mallik, P.K., Fiber Reinforced Composites, New York Marcel Dekker, 1968. Fiberite Manufacturing Procedures, Toolrite Tooling Materials System. United States Gypsum, Tooling Techniques. Morena, J.J., Advanced Composite Mold Making, New York: Van Nostrand Reinhold.
26
CONTROL
26.1 INTRODUCTION Consolidation is an important step that occurs in almost every process used to make an advanced composite article. Consolidation is usually brought about by the application of pressure at a boundary which squeezesair and resin out of the composite thereby changing both its microstructure and dimensions. Improper consolidation can lead to voids, residual stresses,warping and other unwanted effects which could ultimately lead to the rejection of the part. A comprehensive discussion of consolidation in composites would include many complex phenomena. Simultaneously, there is heat, momentum and mass transfer, accompanied by the chemical curing reaction of the resin and the deformation and motion of fibers. Consolidation techniques have been used in the fabrication of both thermoset and thermoplastic composite parts, but are more crucial steps in thermoset composite processing. The traditional composite manufacturing process for aerospaceindustry products usually starts with the B-stage impregnated prepregs consisting of fiber preforms and staged resin matrix. Usually the resin content is relatively high. In order to achieve the required composite material properties which are dominated by the fibers, consolidation is used as an important processing step. Figure 26.1 shows a setup of the prepreg
Handbook of Composites. Edited by S.T. Peters.Published in 1998by O1apman & Hall, London. ISBN 0412 540207
Fig. 26.1 Schematic of the prepreg lay-up used in autoclave cure (Springer, 1986).
lay-up used in autoclave cure. In a traditional lay-up process, prepregs with different fiber orientation and architecture are placed in certain order forming a near-net-shape composite structure. During the lay-up operation, whether it is done manually or using a robot, the trapping of air pockets within the structure is unavoidable. Thus a consolidation step after the lay-up operation is necessary.Prepregs are usually provided with relatively low fiber volume fraction. With the consolidation step, the fiber volume fraction of the composite product can be increased and excessive resin can be removed. The basic mechanisms involved in a consolidation process are the fiber deformation and resin flow, which are coupled with thermal effects and the resin cure reaction. A similar consolidation process can also be seen in soil mechanics such as the settlement of a foundation. However, the deformation behavior of fibrous materials is substantially different from that of granular structures and resin flow behavior is strongly affected by thermal effects
Introduction 577 and chemical reaction of the resin. Thus, the study of the consolidation process of fibrous composite materials involves many disciplines. To effectively control a consolidation process, the selection of the equipment and tooling materials is crucial. Major process parameters for a consolidation process include pressure and temperature and both are functions of time and are usually set as operation cycles. Thus the system setup should be able to effectively control the pressure and temperature profile and transfer heat and pressure to composite parts. Figure 26.2 illustrates the process variables applied during autoclave consolidation and cure.
surface roughness. Metals are widely used as tooling materials for composite processing. However, their heavy weight and high cost of machining become disadvantages when complex geometry is involved. Composite tooling materials have been used as alternatives in various consolidation processes. Another tooling component for the consolidation process is the bleeder, which is usually a nonstructural layer of porous cloth or paper which allows the escape or bleed out of excessive gas and resin during the consolidation process. Sometimes the process is called migration. The bleeder cloth or paper is removed after the curing process and is not part of the final composite. Breather material is used to provide a vacTEMPER ATU RE PRESSURE uum path over the surface of the part. Typical materials are glass and mat. They can be stretched over the part contours to ensure an effective vacuum path and sometime also to provide a cushion effect to matched metal tools. Bagging and sealing are crucial to the quality of the composite parts. General requirements for the bag are: (1)the bag must apply curing pressure uniformly; (2) the bag must not leak under molding conditions; and (3) a good vacuum path must be provided in bagging. Silicone rubber vacuum bags are Fig. 26.2 Illustration of the process variables (tem- widely used because of their long service life. perature, pressure) applied during autoclave Moreover, they are repairable and self-healing consolidation and cure (Springer,1986). with respect to pinholes. The initial cost of fabrication is relatively higher. Nylon is an One of the commonly used facilities is an auto- alternative bag material for up to 193°C clave, which is a closed pressure vessel with (380°F) and is usually discarded after use. means for heating and applying pressure and The commonly used form of resin matrix vacuum to its contents. The dimensions of the prepreg has a resin content beyond 40% and composite parts are limited by the size of auto- requires a significant amount of resin bleedout claves. Thus, for large size composite during cure to achieve a cured laminate resin structures, alternative processing techniques content of 28-32%0.Low resin content prepregs have been used, such as vacuum bag molding. have been developed which can be used withIn addition to the equipment, tooling mate- out resin bleedout processes. Since there is no rial has direct influence on the composite part bleedout process, less resin and less bleeder surface quality, dimensional accuracy and material are needed for a consolidation and residual stress. The main considerations for cure process. However, the removal of tooling material include strength, stiffness, entrapped air becomes a more critical aspect of thermal expansion coefficient, hardness and process control.
t
t
i
t
t
578 Consolidation techniques and cure control portion of the load, then Gutowski’s and Kardos’ models are applicable. In the following discussions, both As composite applications were expanded rapidly in the late 1970s and early 1980s, stud- Springer’s model and Gutowski’s model will ies on the process science of composite be presented. Kardos’ model is equivalent to materials became very active, especially in the Gutowski’s model but different process variareas of consolidation and cure (Lindt, 1982, ables are used in the modeling. 1986; Springer, 1982,1986; Loos and Springer, 1983a,b; Halpin, Kardos and Dudukovic, 1983; 26.2.1 RESIN FLOW Loos and Freeman, 1985; Gutowski et al., 1987a,b; Gutowski and Cai, 1988; Dave, The problem of resin flow in composite proKardos and Dudukovic, 1987a,b; Tang, Lee cessing can be treated as flow through fibrous and Springer, 1987; Batch and Macosko, 1988; porous media. In general it can be handled by Kim et al., 1988 and 1989; Connor et al., 1993). Darcy’s law which states that the flow rate is The purpose of these studies has been to find proportional to the pressure gradient applied out the most suitable process parameters and and is related to the porous medium permethen through scientific process modeling to ability and fluid viscosity. The general form of achieve the optimized composite product Darcy’s law in a one-dimensional case is: quality. (26.1) Pioneering work in composite consolidation and cure modeling was led by Springer (Springer, 1982; Loos and Springer, 1983a). where q is the average resin flow rate, K is the Laminated composite structure was consid- preform permeability with the units of length ered with bleeder layers placed on top of the squared, p is the resin viscosity and dp/dx is composite laminates. When pressure was the imposed pressure gradient. In the case of applied transversely to the laminate plate, resin flow into the bleeder, the consolidation excessive resin material which was in a fluid pressure is established between the advancing flow front and the tool or mold surface. state was squeezed out from the laminate. The main issue involved in using Darcy’s Similar consolidation models have also been developed by Gutowski’s group and law in the consolidation process is that neither Kardos’ group (Gutowski, Morigaki and Cai the fiber preform permeability nor the resin vis1987a; Gutowski et al., 1987b; Dave, Kardos cosity is constant over the process. The preform and Dudukovic, 1987a,b). In these models, permeability is a function of the porosity or both the fiber material deformation which is fiber volume fraction, fiber diameter, fiber orihighly nonlinear and the outgoing flow of the entation and fiber architecture. Among them the fiber volume fraction changes substantially resin are considered. As can be seen from the discussion pre- during a consolidation process. Resin viscosity sented later, both approaches are valid within is related to temperature, the cure status and the ranges of parameters considered. cure time and changes dramatically in the Experimental verification results show good process. Usually at the start of a cure process, agreement with these model predictions resin is in the semi-solid state. With the rise of (Gutowski et al., 1987b; Kim et al., 1988; Cai temperature, it becomes fluid. As the degree of and Gutowski, 1989). When resin content is cure increases, it gels and becomes solid. The permeability of fibrous preforms has relatively high and fiber-to-fiber contact is insignificant, Springer’s model can be applied. been studied both analytically and experimenOn the other hand, if fiber volume fraction is tally (Williams, Morris and Ennis, 1974; relatively high and fibers carry a substantial Gutowski et al., 1987b; Lam and Kardos, 1988, 26.2 CONSOLIDATION MODELS
Consolidation models 579 1989; Van Den Brekel and De Long, 1989) using the well-known Kozeny-Carman equation. The estimation formula using the fiber structural variables can be written as:
K =
r; (1- VJ3 4k0 v;
-~
(0.002in/min.carn oill
x 20 ply sample (0.005 in/rnin.com oill
(26.2)
where rf is the fiber radius, V, is the fiber volume fraction so that (1- VJ is the porosity and k, is an empirical constant, called the Kozeny constant, which is usually determined experimentally. For different textile architecture and orientation, the value of k, will be different. Reported experimental data show that for an aligned fiber bundle, k,,= 0.5-0.7 for the longitudinal flow and k, = 11.0 for the transverse flow. For f 45" cross plies, k, = 2.70. For woven type textile preforms, ko = 5.5. It should be pointed out that many experimental results have been reported and the variation of the Kozeny constant in some cases is significant. Also in the transverse direction, a modified Kozeny-Carman equation has been proposed to account for the stop-flow phenomenon when fiber volume fraction reaches the maximum packing efficiency (Gutowski ef al., 198%). Figure 26.3 shows a comparison of measured axial permeability values for aligned fibers with the Kozeny-Carman equation. Resin viscosity can be expressed as an empirical function of temperature and degree of cure (Lee, Loos and Springer, 1982). The expression can be written as: p = p-exp (U/RT+Ka)
0 20 ply somple
Carman- Kozeny Eq, kxx=0.7 x 5
Fiber Volume Fraction
(Vf)
Fig. 26.3 Comparison of measured axial permeability values for aligned AS-4 fibers with Carman-Kozeny equation (Gutowski et al., 198%).
(26.3)
where p- is a constant, U is the activation energy for viscosity, a is the degree of cure and K is a constant which is independent of temperature. Experimental study has been performed for the Hercules 3501-6 epoxy resin which is widely used in composite fabrication. Figure 26.4 shows the viscosity measurement as functions of temperature and time. To match the model predictions and experimental data, the constant K is found by fitting a linear least square curve to the p versus a data generated
Fig. 26.4 Measured viscosity of 3501-6 resin as a function of time (Lee et al., 1982).
at a constant temperature. Thus the value of K is found to be 14.1 1.2. The values of p, and U are found to bep- = 7.93 x Pa s, U = 9.08 x 104J mol-'. The degree of cure a and the rate of degree of cure da/dt were determined from the results of 'isothermal scanning experiments.
*
580 Consolidation techniques and cure control When the permeability and the resin viscosity are known, with the imposed applied pressure condition, the rate of the outgoing resin flow can be calculated using the Darcy equation. In general, flow may be multi-direcd a l d t = ( K , + K,a) (1- a ) ( B - a) (26.4) tional. Thus 2-D or 3-D flow equations have to be solved. In practice, resin flow in one partica 10.3 ular direction may be dominant, and the (26.5) analysis can be handled as 1-D permeable da/dt = K3 (1- a ) flow. a > 0.3
Efforts were made to describe the d a l d t versus a data with a modified Arrhenius type equation. The proposed empirical equations are
where
26.2.2 FIBER DEFORMATION
K , = A, exp (-AE,/RT)
AE, = 5.66 x lo4J mol-'
The main contribution from Gutowski's model is the description of fiber deformation behavior. Instead of treating fibers as separate layers, a network concept is introduced. In other words, fiber-to-fiber contact is assumed within a fiber assembly, even in the case of aligned fiber bundles. Thus a fiber filament span between the neighboring contact points becomes a small bending beam. During a consolidation process when fibers are pushed closer, more and more fiber-to-fiber contacts take place, and the span length reduces. Thus the bending stiffness of these small fiber beams increases rapidly, resulting in nonlinear elastic deformation response. The nonlinear elastic response of a fiber assembly under a compressive load has been also studied in the textile field, and an empirical formula was proposed (van Wyk, 1946). A proposed fiber deformation model for aligned fiber bundles considers the deformation status variable, the fiber volume fraction V f ,as a function of the consolidation pressure (Gutowski, 1985).The expression is
As can be seen from the discussion, all the constants involved in the model are determined experimentally through a specified process. Similar treatment can be used for other types of resin systems, and experimental investigation results have been reported, including Hercules HBRF-55 Resin (Bhiet aI., 1987)and Fiberite 976 Resin (Dusiet al., 1987).A similar process model has also been discussed by Roylance (1988).
where V , is the maximum obtainable fiber volume fraction for a given fiber network configuration, and V , is the fiber volume fraction below which the fiber network carries no load. The empirical constant As is obtained from curve fitting on available measurement data. A typical fiber deformation curve for
K2 = A, exp (-AE,/RT)
K, = A, exp (-AE,/RT) A,, A, and A, are the pre-exponential factors, AE,, AE, and AE, are the activation energies, R is the universal gas constant, and T is the absolute temperature. The constants in the expression are found as:
B = 0.47 A, = 2.101 x lo9 min-'
A, = - 2 . 0 1 4 ~ l O ~ r n i n - ~ A, = 1.960 x lo5 min-l AE, = 8.07 x lo4J mol-' AE, = 7.78 x 104 J mol-'
Consolidation models 581 well aligned graphite fibers is shown in Fig. 26.5 with the- co-mparison of measured data points.
In
,Q 700
b
In v)
? 500
Data Point-
al
-1 300 e
:200
.c
100
n-
0.4 0.5 0.6 0.7 0.8 0.9 Fiber Volume Fraction ( V f )
Fig. 26.5 Typical fiber deformation curve for wellaligned XA-S and A S 4 graphite fibers (Gutowski et al., 198%).
26.2.3 CONSOLIDATIONMODELS
As discussed above, in Springer’s model, it is assumed that there is no fiber-to-fiber contact. Thus a dynamic fluid pressure exists between the consolidated layers. The consolidation time, which is crucial to the cure process, is related to the permeability of the fibrous preforms, resin viscosity, and the applied consolidation pressure. In Gutowski’s model, the fiber reinforcement and the fluid state resin are considered as a system. Both fiber network deformation and fluid resin flow are solved together. Both models are presented here with a laminated composite structure as the example. The example for Springer’s model is the laminate consolidation with flow in the laminate transverse direction, or z direction. A bleeder ply is assumed to be placed on top of the composite. Figure 26.6 shows the setup for the model. At any instant of time the liquid velocities in the bleeder Vb and in the composite V care given by Darcy’s law. For a constant viscosity liquid, the integrated forms are:
K c (Po - P,)
The proposed relationship between the compressive fiber stress ofand fiber volume fraction V , provides a tool to estimate the finished consolidation status of the composite products. If the time window for the consolidation is long enough, and excessive resin is completely squeezed out from the structure, the consolidation pressure is then balanced by the fiber stress. However, because of the dramatic change of the resin viscosity and preform permeability during a consolidation process, resin flow may not be complete. Thus, developed consolidation simulation models are needed for the process analysis and improvement. During the compression of fibrous preforms, structural relaxation has been observed (Gutowski, 198%). Thus the deformation to some extent is not elastic but viscoelastic. This issue has been addressed by using a Maxwell type model (Kim, McCarthy and Fanucci, 1991).
9, =
7
(26.7)
hC
where p , and pb are the pressures at the composite-bleeder interface and in the bleeder respectively, po is the consolidation pressure and is related to the applied force or pressure, is the instantaneous thickness of the liquid in the bleeder, and hc is the thickness of the resin starved layer, or the thickness of the layers through which resin flow takes place, and K, and Kb are the permeability of the composite layer and bleeder respectively. If the compacted composite layer thickness is h,, then hc = nh,
(26.9)
where n is the number of layers or plies already compacted.
582 Consolidation techniques and cure control
L
Resin Flow
Fig. 26.6 Illustration of the consolidationmodel proposed by Springer (1982).
layers. The final status of the composite is dependent on the compaction of each individual layer. -~ d(hA) = Aq, = Aq, As a comparison, Gutowski’s consolidation (26.10) dt model combines the flow of resin through where A is the surface area of the composite porous media and the fiber deformation laminate, and h is the total thickness of the behavior. Similar treatment has been precomposite laminate. The second equation sented in studies of other fields including soil expresses the fact that at any instant of time, mechanics (Biot, 1941,1955,1956; Gibson and the flow out of the composite is equal to the Hussey, 1967).In general, consolidation occurs flow into the bleeder. The pressure po is related in only one direction, but flow may take place in all three directions. Thus an element is to the applied force as: deformable in the z direction. A new variable 6 f (26.11) is used to represent the deformation, and 6 = z = + pa + w where w is the local displacement of the where F is the applied force and pa is the fiber network. The laminate setup for the atmospheric pressure. By combining these model is illustrated in Fig. 26.7. If the initial equations, the consolidation equation fiber volume fraction for the composite is V, and the fiber volume fraction at any instant is becomes: Vf, the fiber continuity condition states The equation of continuity gives the rate of change of volume of the composite as:
vo=-v, 36 Therefore for each individual layer, the consolidation time can be calculated. The total consolidation time is the summation for these
aZ
(26.13)
Resin flow continuity condition requires:
Consolidation models 583 Here it is assumed that the inertial effects in the process are small. Therefore the applied pressure is balanced by a combination of the average resin pressure and the fiber stress. In (26.14) other words, any load which is carried by the fibers is then unavailable for pressurizing the With the application of Darcy's law, a consoli- resin. dation equation using the fluid pressure p , and Since both the permeability and the fiber fiber volume fraction Vf as variables can be stress are expressed as functions of fiber volwritten as ume fraction V , with the given initial and boundary conditions, the variables V ,and p , as a function of time and location can be solved. In general numerical calculation procedures have to be developed for solving the partial differential equations. In some simplified cases, analytical solutions are possible. This equation gives a relationship between the spatial and time-varying nature of the presExample problem 1: One-dimensional flow sure in the resin and the fiber volume fraction in compression molding of the composite. The equilibrium statement for the consolidation is: A simplified example of composite consolidation is the compression molding of a flat (26.16) rectangular laminate. The composite part is A pressed between two solid dies. Therefore only in-plane flow is possible. In other words, flow components are in the x and y directions only. If the initial fiber volume fraction is uniform, the equation of the resin flow and fiber deformation becomes:
J2Pr + K - a2pr + -~ P av, = 0 (26.17) K
X
~y
ay2
v,
at
ho Here it is also assumed that there is no signifi-
a€
Fig. 26.7 Illustration of the consolidation model proposed by Gutowski et al. (1987a).
cant pressure gradient in the z direction, and the viscosity p does not vary spatially. In some cases, K Z / a 2>> Ky/b2 where a and b are the dimensions of the laminate in x and y directions respectively. The compression molding results in primarily one-dimensional flow in the x direction. Then the equation can be solved analytically. With the assumed boundary conditions of p , = 0 at x = M and ap,/ax = 0 at x = 0, the result is a parabolic pressure distribution as
584 Consolidation techniques and cure control
The solution for the fiber volume fraction Vf as a function of time is:
Po = Of(Vf) 3 pa' K _ Vdvf ,T +
*
(26.19)
L
This expression shows how the applied pressure p, is carried by the fiber stress G~ and the average pressure in the resin. The load sharing in a composite is directly analogous to how the load is shared in a parallel spring and damper set. For example, initially if Vf is less than V,, then there is no deformation in the spring (fibers) and the entire load is carried by the resin. On the other hand, at long times and finite viscosity, if the rate of change of Vf is close to zero, then the pressure in the damper (resin) goes to zero and the total load must be carried by the fibers. Figure 26.8 shows an example of the one-dimensional flow in compression molding with the comparison of computer simulation results.
Example problem 2 Compression molding with two-dimensional flow Here the case of compression molding of a rectangular laminate with an isotropic in-plane permeability is considered. In other words, Kx = Ky = K. This may correspond to a quasiisotropic lay-up. The flow equation becomes Poisson's equation, which can be solved by the separation of variables technique. The solution for the pressure distribution in a laminate with zero pressure at the boundaries is:
With the applied load balance condition, the final result is:
It can be seen that the result is analogous to the previous case except for a geometry effect term which is shown in the bracket. 600
- PR,Theory
- 400 -
0
PR,Measured
'3 0
Time ( m i d
Fig. 26.8 Example of one-dimensional flow in compression molding and computer simulation results (Gutowski, et al., 198%).
Example problem 3: Bleeder ply molding This has been presented with the Springer's model. In this case, a porous bleeder ply is placed on top of the composite, and flow is principally in the z direction. With the introduction of a new variable, the void ratio e = (1- Vf)/V, one may obtain the nonlinear one-dimensional consolidation equation. An equation similar to this was first derived by Gibson et al. (1967) for the consolidation of saturated clays. The expression is:
(
")
de = (e, + 1)2- a - Kz 'Of at 3.z p ( l + e ) ' e az
-
(26.22)
Consolidation models 585 The void ratio e or the fiber volume fraction V, is a function of both time and location. An equivalent equation using variables Vf and p , can be written as
With similar pressure equilibrium conditions, the distribution and time history of Vf or e can be solved numerically. Figure 26.9 shows an example of the bleeder ply molding measurement setup, and the comparison of the computer simulation results with the measured data. It is interesting to see that, with Gutowski’s consolidation model, the final status of the composite in terms of the average fiber volume fraction can be estimated from the proposed fiber deformation model if the consolidation process is complete. The consolidation time for a particular setup can be solved through numerical simulation. As can be seen from the analysis, the total consolidation time for a composite structure is strongly dependent on the dimension in the resin flow direction. For laminated composite structures, usually the dimensions in x and y directions (directions within the laminate structure) are much larger than that in the z direction (direction transverse to the laminate plane). For example, many aerospace structural parts range from a few inches to several feet in x or y direction, but only have a thickness of a fraction of an inch in the z direction. Thus the bleeder ply molding process is preferred and is widely used in many part fabrication processes. However, for the socalled thick composites, for example with lay-up of 64 or 96 plies, the consolidation time required increases dramatically in the bleeder ply molding cases. With the selected cure cycle for thin composites, complete consolidation may not be achieved for thick composites. Thus the final fiber volume fraction of the thick composite tends to be relatively lower. This has been observed in experiments involving thick
DT TRANSDUCER
600
/Applied
Pressure
doto -theory o
500 -
-.-
Modified Cormon-Kozemy
400-
v)
a v 300
-
3
L
a
\o
IO0
\O
P 0 0
10
20
30
40
50
60
Time (rnin) Fig. 26.9 Example of bleeder molding and computer simulation results (Gutowskiet al., 198%).
composites (Kim, Jun and Lee, 1989). It can also be seen from the comparison of the two models that with relatively low fiber volume fraction, fibers carry almost no load. Thus the consolidation process is dominated by the resin flow through the fiber network. Then the difference between the two models is very minor. Springer assumes the consolidation is done layer by layer, while Gutowski treats the fiber network as a whole system. However, in both cases the top layers are consolidated first. When the fiber volume fraction becomes high, then the predictions from the
586 Consolidation techniques and cure control
two models show significant different results. On the other hand, the numerical schemes of the two models are different. Springer’s model requires only the solutions of a series algebraic equations, while in Gutowski’s model nonlinear partial differential equations have to be solved. A comparison study has been presented by Smith and Poursartip (1993). 26.3 CURE CONTROL
Fiber reinforced thermosetting resin composites manufactured in autoclaves are made by forming the uncured fiber-resin mixture into the desired shape and then curing the material. Curing requires the application of heat and pressure. Heat is used to facilitate and control the chemical reactions of the resin, and pressure is used to consolidate the composite, squeeze out the excess resin, and minimize the void content. A cure cycle usually means the magnitude, duration, and profile of the temperature and pressure applied during a curing process. Selection of the cure cycle directly affects the quality of the finished composite product, such as fiber content, fiber distribution, and void percentage.
Viscosity
Specifically, a selected cure cycle must ensure that:
the temperature inside the material does not exceed a preset value at any time during the cure; 2. at the end of the cure the resin content is uniform and has the desired value; 3. the material is cured uniformly and completely; 4. the cured composite has the lowest possible void content; 5. the cured composite has the desired thermal and mechanical properties; 6. the curing is achieved in the shortest time. Figure 26.10 shows schematically the overall cure process model structure. In an early study, Loos and Springer (1983a)proposed a thermochemical model. Heat transfer from the environment to the composite material determines the temperature distribution, the degree of cure of the resin, and the resin viscosity within the composite structure. The temperature inside the composite can be calculated using the law of conservation of energy. By neglecting the energy transfer by convection, the energy equation can be expressed as:
b
Flow
/
Reaction kinetics
\
Heat transfer
Residual stress
Fig. 26.10 Schematic of overall cure process model (Dave et al., 1990).
Cure control 587 resin viscosity, the degree of cure a and the rate of the cure da/dt can be characterized using a modified Arrhenius type equation, with relevant constants in the model deterdH -k+p(26.24) mined experimentally (Lee, Loos and dt Springer, 1982; Bhi et al., 1987; Dusi et al., 1987; where p and Cv are the density and specific Roylance, 1988).Figure 26.11 show an example heat of the composite, kx, k and k, are the ther- of the rate of heat generation and rate of mal conductivities, and ?is temperature. In degree of cure of the 3501-6 resin system as the case of relatively thin composite structure, functions of time and temperature. conduction heat transfer is mainly in the z It is noted that the densityp, specific heat Cy, direction. Thus terms in the x and y directions heat of reaction Hr, and thermal conductivity k can be dropped. The rate of heat generation are all dependent on the instantaneous and dH/dt is defined as: local resin and fiber contents of each ply, and
i ~zE) (
(26.25) 01
c3501-6
400K
1 05
450K
1
where H, is the total heat of reaction depending on the resin type. The rate of the cure reaction is a function of temperature and the cure status, and can be expressed symbolically as: da dt
- = f(T,
4
(26.26)
The degree of cure is then determined as:
)$(:I
(26.27)
dt
a =
o f \
It is assumed that for an uncured material, a = 0, and for a completely cured material, a approaches unity. As discussed earlier the
0 0
,
02
,
Iojo\J
06 0 02 DEGREE OF C U R E , a
04
04
06
TEMPERATURE ( K l I
I
1
1
I
I
3501 - 6
-
-1
0 5 r n coI/sec
-
s \ I
,
,
,
,
~
,
,
l
,
l
l
l
l
Fig. 26.11 Rate of heat generation and rate of degree of cure of the
588 Consolidation techniques and cure control can be handled using rule of mixtures (Loos and Springer, 1983c) or proposed approximate formulas (Springer and Tsai, 1967). The solution to these equations can be obtained once the initial and boundary conditions are specified. The initial conditions require that the temperature and degree of cure inside the composite be given before the start of the cure. The boundary condition requires that the temperatures on composite surfaces in contact with the tool be known as a function of time during cure. Therefore the boundary condition is related to the specified cure cycle and the equipment setup. The objective for the cure control scheme is to achieve the desired composite quality. Some of the main targets are reasonable temperature distribution, complete consolidation, minimum thermal stress and minimum void content. With a developed numerical scheme, the temperature distribution inside the laminate is calculated as a function of position and time. A good cure scheme should realize the two main targets: (a) the temperature is reasonably uniform inside the material and (b) the temperature does not exceed a preselected maximum at any time. For a given cure temperature and cure pressure, the time window for the consolidation is then specified. From the consolidation models, the compaction status of the consolidated composite can be obtained. In Springer’s model, the result is the total number of compacted plies, while in Gutowski’s model the result is the V ,distribution across the layers. If the consolidation cannot be completed with the selected cure cycle, proper modifications are then made. The compaction issue becomes crucial to the cure process of the thick composite structure. A multiple stage heating process may be designed to defer the cure reaction of the resin and thus prolong the consolidation time window. Voids within the composite material are harmful to its mechanical Performance. Experimental study shows that the resin
pressure early in the cure cycle and the initial resin moisture are crucial considerations in producing void-free laminates (Kardos et al., 1983, 1988). Since the driving force for diffusion rises with temperature, in order to prevent the potential for pure water void growth by moisture diffusion in a laminate at all times and temperatures during the curing cycle, the resin pressure at any point within the curing laminate must be higher than the minimum resin pressure required, which is a function of the relative humidity and temperature (Dave et al., 1990).Figure 26.12 shows a void stability map for pure water void formation in epoxy matrices. A similar pressure requirement also holds for small air/water voids after an initial growth period. It has also been observed that the void content is reduced (1 ATM I 101 kPI)
(RH), = 1ooO/o
300
400 1,K
(RH),= 50%
500
Fig. 26.12 Void stability map for pure water void formation in epoxy matrices (Dave et al., 1990).
Efects of tooling and part shape 589 significantlywhen the applied pressure is sufficiently high to collapse the vapor bubble before the gel point is reached. Therefore, after the time-temperature cycle is determined, it is possible to obtain a profile of the minimum pressure versus cure time. The boundary pressure is then maintained greater than the minimum pressure throughout the cure cycle. During the cooling stage after the cure of the composites, residual thermal stress is related to the difference between the cure temperature and ambient temperature, and the thermal expansion behavior of the composite material. For a laminated structure, calculation of the thermal stress has been discussed and formulated by Tsai and Hahn (1980). Since the material shows viscoelastic behavior, stress relaxation has been observed over time. A post-cure process is usually applied to the structure to relieve the induced thermal stress. For large complex-shaped composite structures, non-autoclave curing methods are used. Compared with traditional autoclave curing methods, the component size restrictions are eliminated, energy consumption is reduced, and capital equipment cost can be cut down. The non-autoclave processes use an oven, integrally reinforced tools, and presses. Major issues related to non-autoclave curing are the effective compaction of the composite plies, and the elimination of the trapped interlaminar or intraply air. 26.4 EFFECTS OF TOOLING AND PART SHAPE
Properly designed tools that produce acceptable parts on a reproducible basis are a must when fabricating composite structures. The tool design requires the consideration of as many factors as are studied in the design of the part itself. The main requirement for the tools is to maintain proper geometric dimensional stability and surface profile during the compression and thermal cycling processes. On the other hand, the tool must also be
heated to a specified temperature at a specified rate under controlled conditions in the autoclave. Tooling materials may be metal (steel, nickel, nickel alloys, and aluminum), graphite-epoxy and elastomer, depending on different composite part shape, size, volume of production and curing method. Selection of the tooling material often reflects a compromise among these considerations. Thermal behavior of the tooling material is also crucial in the design and fabrication. Table 26.1 lists the coefficient of thermal expansion of different composite and tooling materials. The values for the composites are dependent on the ply orientation and fiber volume fraction, and typical values are shown there.
Table 26.1 Coefficient of thermal expansion (CTE) for various materials (Borstell and Turner, 1987)
Material
CTE ( I P / K )
Structural composite material Boron-poxy Aramid+poxy Graphiteepoxy Fiberglass-epoxy
3.6-10.8 -2.0-5.8 1.8-9.0 7.2-9.0
Tooling material Graphiteepoxy Cast ceramic Tool steel Iron (electroformed) Nickel (electroformed) High-temperature cast epoxy Aluminum Silicone rubber
4.1-9.0 0.81 11.3 11.9 12.6 19.8 23.2 81-360
26.4.1 TOOLING FOR AUTOCLAVE MOLDING
The traditional autoclave molding process uses a vacuum bag to impose a pressure difference on the composite lay-up. A typical bagging system consists of the following steps (Schwartz, 1983).
590 Consolidation techniques and cure control 1. Cover the lay-up with a perforated parting film or separator. Then lay up a layer or layers of bleeder material. The requirement of the bleeder layers should be such as to ensure adequate bleeding of air and excess resin out of the part. 2. Place a strip of jute (vent material) just beyond the edge of the lay-up and put bagsealing compound along the outside perimeter. 3. Cover the lay-up, jute, and sealing compound with a flexible-film diaphragm and seal the diaphragm to the mold with the seal compound. 4. Connect the vacuum lines and slowly apply the vacuum pressure while working the wrinkles and excess air out of the lay-up, bleeder material, and vacuum bag. 5. Check system for vacuum leaks. 6. Keep the part under vacuum while it is waiting to be cured in the oven or autoclave.
Graphite-epoxy laminate
Mold form
-/
cdp\
Mold half-
Angle caul plate
Caul plate stop
{Resin reservoir
.Mold half
To prevent surface irregularities on the bag side (untooled surface) of the parts, a caul plate may be used. The sole purpose of caul plates is to improve the visual appearance of I the parts. They do not control part thickness. A flexible caul plate with a thermally stable rub\Cao ber such as silicone or a fluoroelastomer is often used to accommodate the surface geom- Fig. 26.13 Example of autoclave tooling (Borstell and Turner, 1987). etry. Figure 26.13 shows examples of autoclave tooling setups with caul plates. The three issues related to the tooling introducing a thermal strain. As the part and design (Borstell and Turner, 1987) are thermal tool cool down from the gel temperature, the expansion correction, coordinating the loca- tool usually shrinks more than the part. As an alternative, graphiteepoxy molds are used in tion of partial plies and use of caul plates. Because of the low coefficients of thermal some applications. Although some data has expansion of composites when compared with been published, not all composite materials metal tooling materials, thermal strain or have been measured. One empirical method is stress must be considered for a curing process. to cure a representative panel on a plate of the In the autoclave, the temperature at which the specified tooling material using the specified resin solidifies is the gel temperature. At that cure cycle. Corrections can be estimated by specific temperature, the part is the same size comparing the difference between the mold as the thermally expanded mold. At a temper- and part dimensions. Another recommended ature above the gel temperature, the tool empirical correction method is to correct steel expands more than the partially cured part or nickel tools by making the tool 0.999 of the
Effects of tooling and part shape 591 engineering dimension, and to correct aluminum tools by 0.998. For example, a 2540 mm (100 in) dimension is tooled to be 2537 mm (99.9 in) for the steel tool. These corrections are needed to ensure an acceptable fit of mating composite parts. Most parts contain partial plies to accommodate local areas of increased stress. Several techniques are used to control the location of partial plies, including polyester film templates, slotted templates, and rails and banking surfaces. These tools serve as supplemental guidance to position the partial plies in the lay-up process. Typical cases of applying a caul plate are to control the edge of a panel or the flanges of channels. The design of the metal caul plates must take into account the fact that the matrix resin melts in the autoclave to a very low viscosity. The caul plate performs by pushing excess resin sideways. Thus the rigid metal caul plates must have high rigidity so that they do not deflect under autoclave pressure at curing temperature. The thickness of the caul plates can be calculated by use of the equations for unsupported bending beam analysis. The deflection of the caul plate can be estimated using the balance condition of resin pressure and applied force (Gutowski and Cai, 1988).The caul plate deflections should be limited to half the tolerance permitted in the part.
aluminum. During autoclave curing of composite parts, the thermal uniformity is excellent with rapid heat-up and cool-down rates. It is easy to handle and transport because of its light weight. It also offers outstanding durability because the mold surface resists cutting or impact damage and is not thermally degraded. When damaged, it is easy to repair by welding, soldering, silver-soldering, or selective plating. It can provide complex contours without expensive machining. With most resin systems, it shows good release properties. Figure 26.14 shows the procedures of making an electroformed nickel tool. As in some other types of tooling, constructing a model of the part surface is the first step in creating an electroformed mold. The models are the same net dimensions as the required nickel mold. Compensation may be required when the coefficient of thermal expansion of the composite part differs greatly from that of the nickel mold. Models are made from plaster, epoxyfaced plaster, fiberglass, fiberglass-epoxy, wood or other materials. From the model a reverse mandrel 'splash' is generally fabricated from epoxy-faced fiberglass or plaster. The mandrel to be used in electroforming is then copied from the 'splash', although the model can be used as the mandrel if it is prepared correctly. The comers of the mandrel should be designed to have radii in excess of 0.76 mm (0.030 in) to avoid thin spots in the 26.4.2 ELECTROFORMED NICKEL TOOLING deposit. Draft and taper should be designed An electroformed nickel tool consists of a into the mandrel to facilitate its removal from 4.6-6.4 mm (0.18-0.25 in) thick electrode- the electroform. Sharp corners or narrow, deep posited mold surface that is supported by a grooves should be avoided if possible. The simple steel substructure. The mold surface is mandrel can be fabricated from epoxy-faced produced by the electroplating process fiberglass, rubber, or other materials. The surface of the mandrel is made conductive by (Sheldon, 1987). The electroformed tooling concept offers proper coatings. The back of the mandrel must numerous advantages. The size of the mold is be reinforced to keep the mandrel from disrestricted only by the size of the electroform- torting during the electroforming process. Electroforming is the process of producing ing tank. The cost of producing duplicated tools is low. The mold surface is very smooth an article by electrodeposition of a metal onto and scratch resistant. The coefficient of ther- a conductive mandrel surface. An anode susmal expansion is approximately 40% less than pended in an aqueous electrolyte is connected
592 Consolidation techniques and cure control
---t
Splash
Model
-
Fiberglass plating mandrel
Mold electroformed
I
I
i
Plated mold and tool upport structure
Mold and structure joined
Plating mandrel removed
Fig. 26.14 Example of electroformed nickel tooling (Sheldon, 1987).
to the positive pole of a DC electric source, and parts. These include low coefficient of thermal the mandrel (cathode) is connected to its neg- expansion, ease of preparation, low density, ative pole. The flow of electricity or electrons and thermal stability (Harmon, 1987). Their results in the oxidation of a nickel anode to disadvantage is that they are less durable than nickel ions and the reduction of nickel ions to metal tools. Composite tool making starts with a master nickel metal at the cathode (mandrel). The typical rate of growth is approximately model, usually built with plaster or hard0.013-0.025 mm (0.0005-0.001 in) per hour. wood. The master models require proper When the electroform is removed from the drying, sealing, and coating with mold release. mandrel, its surface is a mirror image of the Then lay-up can be done directly on the plaster surface of the mandrel. A natural physical or wood master. Liquid gel coats are required characteristic of electrodeposition is that elec- to obtain a high fidelity surface on tools cured tric current will tend to localize the deposit on by the vacuum bag process which does not all edges and corners, causing an uneven generate enough pressure to ensure a void-free thickness on the electroform. However, there surface, but may not be required on tools are a variety of techniques to offset this effect. cured by the autoclave process which does After the desired mold thickness is provide sufficient positive pressure. Prepregs obtained, the mold is removed from the tank, with light weight fabrics are used directly cleaned and the steel back-up structure is against the tool surface, while prepregs with attached. The nickel mold is then polished to heavier fabrics are used to build up the thickness. During the lay-up, care should be taken the required finish, and ready for use. to work each ply into all radii and corners and to remove all entrapped air. Debulking is 26.4.3 GRAPHITE-EPOXY TOOLING applied after the lay-up, either with a vacuum Composite tools have definite advantages bag setup or with -assistance of an autoclave over metal molds for large or highly contoured for a pressure debulk, to consolidate the plies
Eflects of tooling and part shape 593 and remove all entrapped air. The curing process is done with a vacuum bagging system or with an autoclave. With the tool still on the model, the support structure, either a solid laminate or an ’egg-crate’ panel is attached to the tool by means of locally applied fabrics, room-temperature curing, and high-temperature resistant resins. Once the support structure is cured to the laminate shell, it is removed from the master. Care should be taken to avoid damaging either the tool or the master. Figure 26.15 illustrates the graphite-epoxy tooling making process. Compositetools are being used successfully throughout the aerospace industry to produce parts that are structurally reliable, reproducible, and dimensionally accurate.
In thermal expansion molding, two basic methods are employed: the trapped or fixedvolume rubber method and the variable-volume rubber method. Figure 26.16 shows the setup for both methods. The fixedvolume method exploits the large difference between the coefficient of thermal expansion of the elastomer and that of metals. The elastomer is confined within a closed metal tool
Rubber tool sized to fill the cavih, in the pan
,Pan Teflon separator film
Breather cloth
/
Vacuum bag
.---
Floating-plate pressure control
-
’
Rubber tool projects above the pan 30 excess pressure is vented by forcing the floating plate to the bag.
‘PFP
master
Fig. 26.15 Example of graphite-epoxy tooling
(Harmon, 1987).
M
,Outer
26.4.4 ELASTOMERIC TOOLING
Elastomeric tooling or rubber tooling can be used to generate molding pressure or to act as a pressure intensifier. In thermal expansion molding, elastomeric tooling is constrained within a rigid frame to generate consolidation pressure by thermal expansion during the curing cycle (Foston and Adams, 1987).
box
w
Fig. 26.16 Example of elastomeric tooling (Foston
and Adams, 1987) (a) fixed volume method; (b) variable volume method.
594 Consolidation techniques and cure control Impregnated Composites, Proc. 9th Int. Cod. cavity. When heated, it expands into the cavity, Composite Mater. (ICCM-9), 1993, 3,575-583. exerting the pressure required to compact a composite laminate. The variable-volume Dave, R.S., Kardos, J.L. and Dudukovic, M.P., A Model for Resin Flow During Composite method offers more flexibility and control than Processing, Part 1: General Mathematical the fixed-volume method because a precisely Development, Poly. Composites, 1987, 8(1), calculated volume of rubber is not normally 29-38. required. In most applications, the rubber is Dave, R.S., Kardos, J.L. and Dudukovic, M.P., A Model for Resin Flow During Composite simply 'set back' to allow for the bulk factor of Processing, Part 2: Numerical Analysis for the molding material during assembly of the Unidirectional Graphite/Epoxy Laminates, tooling details. A floating plate is used for the Poly. Composites, 1987,8(2), 123-132. pressure control. Dave, R.S., Mallow, A., Kardos, J.L. and Dudukovic, Thermal expansion molding with elasM.P., Science-based Guidelines for the tomeric tooling has been successfully used on Autoclave Process for Composites commercial aircraft parts such as rudders and Manufacturing, SAMPE I., 1990,26(3),31-38. spoilers (Schneider and Carroll, 1987). This Dusi, M.R., Lee, W.I., Ciriscioli, P.R., and Springer, G.S., Cure Kinetics and Viscosity of Fiberite 976 reduces the number of detail parts fabricated Resin, J. Composite Mater., 1987,21(3),243-261. and the need for bonding and mechanical fasFoston, M. and Adams, R.C., Elastomeric Tooling, tening on assembly, thereby effecting in Engineered Materials Handbook, Vol. 1: significant reductions in production time and Composites, ASM International, 1987, pp. cost. 590-594. Gibson, R.E. and Hussey, M.J.L., The Theory of One-Dimensional Consolidation of Saturated REFERENCES Clays, Geotechnique, 1967,17,261-273. Batch, G.L. and Macosko, C.W., A Model for Two- Gutowski, T.G., A Resin Flow /Fiber Deformation Model for Composites, S A M P E Quarterly, 1985, Stage Fiber Deformation in Composite 16(4),58-64. Processing, Proc. 20th Intern. SAMPE Tech. Gutowski, T.G., Morigaki, T. and Cai, Z., The Conf., September 1988, pp. 641-650. Consolidation of Laminate Composites, J. Bhi, S.T., Hansen, R.S., Wilson, B.A., Calius, E.P., Composite Mater., 1987,21, 172-188. and Springer, G.S., Degree of Cure and Viscosity of Hercules HBRF-55 Resin, Proc. 32nd Intern. Gutowski, T.G., Cai, Z., Bauer, S., Boucher, D., Kingery, J. and Wineman, S., Consolidation SAMPE Symp. Exhib., Vol. 32., 1987, pp. Experiments for Laminate Composites, J. 1114-1118. Composite Mater., 1987,21,650-669. Biot, M.A., General Theory of Three-Dimensional Gutowski, T.G. and Cai, Z., The Consolidation of Consolidation, J. Appl. Pkys., 1941,12, 155-164. Composites, in The Manufacturing Science of Biot, M.A., Theory of Elasticity and Consolidation Composites, Proc. Manufacturing International for a Porous Anisotropic Solid, J. Appl. Phys., 88, Vol. IV,(ed T.G. Gutowski), 1988, pp.13-25. 1955,26(2), 182-185. Biot, M.A. , General Solutions of the Equations of Halpin, J.C., Kardos, J.L. and Dudukovic, M.P., Elasticity and Consolidation for a Porous Processing Science: An Approach for Prepreg Material, J. Appl. Meck., 1956, March, 91-96. Composite Systems, Pure Appl. Chem., 1983,55(5). Borstell, H. and Turner, K.T., Tooling for Autoclave Harmon, B.D., Graphite-Epoxy Tooling, in Molding, in Engineered Materials Handbook, Vol. Engineered Materials Handbook, Vol. 1: Composites, 1: Composites, ASM International, 1987, pp. ASM International, 1987, pp.586-589. 578-581. Kardos, J.L., Dudukovic, M.P., McKague, E.L. and Lehman, M.W., Void Formation and Transport Cai, Z . and Gutowski, T.G., Fiber Distribution and Resin Flow in the Molding Process, Proc. 7th During Composite Laminate Processing: An Initial Model Framework, in Composite Int. Conf. Composite Mater. (ICCM-7), 1989, 1, Materials: Quality Assurance and Processing, 76-82. ASTM STP 797, (ed C.E. Browning), 1983, pp. Connor, M., Gibson, A.G., Toll, S. and Manson, J.A.E., A Consolidation Model for Powder 96-109.
References 595 Kardos, J.L., Dave, R. and Dudukovic, M.P., Voids in Composites, in The Manufacturing Science of Composites, Proc. Manufacturing International '88, Vol. IV,(ed T.G. Gutowski), 1988, pp. 4148. Kim, T.W., Yoon, K.J., Jun, E.J. and Lee, W.I., Compaction Behavior of Composite Laminates During Cure, SAMPE I., 1988,24 (S), 33-36. Kim, T.W., Jun, E.J. and Lee, W.I., Compaction Behavior of Thick Composite Laminates During Cure, Proc. 34th Inter. SAMPE Symp., 1989, 12 (l),17-19. Kim, Y.R., McCarthy, S.P. and Fanucci, J.P., Compressibility and Relaxation of Fiber Reinforcements During Composite Processing, Polym. Composites, 1991,12 (l),13-19. Lam, R.C. and Kardos, J.L., The Permeability of Aligned and Cross-Plied Fiber Beds During Processing of Continuous Fiber Composites, Proc. Am. SOC.Composites, Third Technical Conf., Seattle, WA, 1988, pp. 3-11. Lam, R.C. and Kardos, J.L., The Permeability and Compressibility of Aligned and Cross-Plied Carbon Fiber Beds During Processing of Composites, Proc. 47th Ann. Tech. Conf. (ANTEC'89), SPE, New York, 1989, pp. 1408-1412. Lee, W.I., Loos, A.C., and Springer, G.S., Heat of Reaction, Degree of Cure, and Viscosity of Hercules 3501-6 Resin, J. Composite Mater., November 1982,16, pp. 510-520. Lee, S.Y. and Springer, G.S., Effects of Cure on the Mechanical Properties of Composites, J. Composite Mater., 1988,22(1), 15-29. Lindt, J.T., Engineering Principles of the Formation of Epoxy Resin Composites, SAMPE Quarterly, October, 1982. Lindt, J.T., Consolidation of Circular Cylinders in a Newtonian Fluid, I. Simple Cubic Configuration, J. Rheology, 1986,30. Loos, A.C. and Freeman, Jr., W.T., Resin Flow During Autoclave Cure of Graphite-Epoxy Composites, High Modulus Fiber Composites in Ground Transportation and High Volume Applications, ASTh4 STP 873, (ed D.W. Wilson), 1985, pp. 119-130. Loos, A.C. and Springer, G.S., Curing of Epoxy Matrix Composites, J. Composite Mater., 1983,17, 135-1 69. Loos, A.C. and Springer, G.S., Calculation of Cure Process Variables During Cure of Graphite/Epoxy Composites, Composite Materials: Quality Assurance and Processing, ASTM STP 797, (Ed. C.E. Browning), 1983, pp.
110-118. Loos, A.C. and Springer, G.S., Curing of Graphite/Epoxy Composites, Air Force Materials Laboratory Report AFWAL-TR-834040, Wright Aeronautical Laboratories, Wright Patterson Air Force Base, Dayton, OH, 1983. Roylance, D., Reaction Kinetics for Thermoset Resins, in The Manufacturing Science of Composites, Proc. Manufacturing International'88, Vol. IV, (ed T.G. Gutowski), 1988, pp. 7-11. Schneider, C.W. and Carroll, H.E., Elastomeric Tooling Application, in Engineered Materials Handbook, Vol. 1: Composites, ASM International, 1987, pp. 595-601. Schwartz, M.M., Composite Materials Handbook, McGraw-Hill, 1983. Sheldon, D.L., Electroformed Nickel Tooling, in Engineered Materials Handbook, Vol. 1: Composites, ASM International, 1987, pp. 582-585. Smith, G.D. and Poursartip, A., Comparison of Two Resin Flow Models for Laminate Processing, J. Composite Mater., 1993,27(17),16951711. Springer, G.S. and Tsai, S.W., Thermal Conductivities of Unidirectional Materials, J. Composite Mater., 1967,1, 166-173. Springer, G.S., Resin Flow during the Cure of Fiber Reinforced Composites, J. Composite Mater., 1982,16,400410. Springer, G.S., Modeling of the Cure Process of Composites, SAMPE J., September/October 1986, pp. 22-27. Tang, J.M., Lee, W.I. and Springer, G.S., Effects of Cure Pressure on Resin Flow, Voids, and Mechanical Properties, J. Composite Mater., 1987, 21,421440. Tsal, S.W. and Hahn, H.T., Introduction to Composite Materials, Technomic Publishing, 1980. Van Den Brekel, L.D., and De Long, E.J., Hydrodynamics in Packed Textile Beds, Textile Research J., August, 1989, pp. 433-440. van Wyk, C.M., Note on the Compressibility of Wool, J. Textile lnst., 1946, 37, T285-T292. Williams, J.G., Morris, C.E.M. and E d s , B.C., Liquid Flow through Aligned Fiber Beds, Polym. Engng Sci., 1974,14 (6), 413-419.
COMPOSITE MACHINING
27
Kent E. Kokkonen and Nitin Potdar
27.1 INTRODUCTION
27.2 CONVENTIONAL MILLING
The processes used to manufacture composite When milling graphite-epoxy with polycrysstructures generally require that trimming and talline diamond (PCD) the chips are formed as other machining operations be performed small particles of powder dust and fumes. The prior to assembly. Machining processes are surface roughness is a function of fiber orienrequired to produce accurate surfaces and tation, cutting direction and the angle between holes to allow precision fitting of components cutting direction and fiber direction. The surinto an assembly. Due to shrinkage during the face may sometimes exhibit many small holes curing stage of the composite structure it is not due to fiber pull out. When taking heavy practicable to place holes in the part during milling cuts there is a greater tendency to the molding stage, therefore milling, cutting, break comers as the tool exits the material so it drilling etc. are considered a post cure opera- is advisable to first machine a step on the edge tion. perpendicular to the final pass. A four fluted Due to the toughness and abrasive nature of end mill will reduce cutting pressure on the modern composites, there is a need for harder laminate and keep it cooler. Climb milling and longer lasting cutting tools. A large data- helps prevent the fibers from separating from base of machining information for various the matrix bond material. high speed steel and carbide cutting tool Advantages of machining composites are: materials exists for machining metal, wood and some thermoplastics. However, much of 0 improved surface finish unless part surface was directly in contact with the mold surthis data cannot be applied to machining modface; ern composites. Modern composites like 0 machined surfaces provide accurate mating graphite-epoxy, aramid-epoxy and carbonsurfaces for parts to be assembled; carbon each have their own machining charac0 eliminates the majority of the problems teristics. Composites are not homogeneous or associated with part shrinkage and insert isotropic, therefore the machining characterismovement during the fabrication processes. tics are dependent on the tool path in relation to the direction of the reinforcing fibers. Tool life factors are: Metals or metal alloys have nearly homogeneous properties throughout the workpiece, 0 PCD end milling cutters will perform sixty to one hundred times longer than carbide; but each material in a composite retains its 0 cutting speed does not have a great effect individual properties. on the flank wear of PCD cutting tools. With increased cutting speeds, the feedrates can be increased and machining time Handbook of Composites. Edited by S.T. Peters. Published decreased; in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Mechanical drilling of composite materials 597
0
0
0
cutting speeds range from 244 surface m/min (800 surface ft/min) to 762 surface m/min (2500 surface ft/min) with PCD end mills; when cutting parallel to the fiber direction, the wear ratio on the cutting tool increases compared with cutting perpendicularly to the fiber direction; surface finish remains below 20Ra [arithmetical average roughness (see IS0 R488)] when cutting with PCD end mills and the flank wear is approximately 0.127 mm (0.005in); the surface finish deteriorates above 150 Ra when cutting with a carbide end mill and the flank wear has reached 0.127 mm (0.005in); roughing feedrates range from 0.23 mm/rev (0.007 in/rev) to 0.38 mm/rev (0.012 in/rev) and finish feedrates range from 0.076 mm/rev (0.002 in/rev) to 0.13 mm/rev (0.005 in/rev); the depth of cuts should range from one quarter to one half of the diameter of the end mill cutter. Depth of cut will vary depending on the rigidity of machine ways, spindle and workholding devices.
The disadvantages associated with milling of composites include controlling the graphite chips (dust particles), confining them to a small area and having an adequate collection system. A second problem is controlling the outer layers of the composite so that the fibers will shear instead of lifting up under the force of the cutting action and leaving extended fibers beyond the cut surface. Also when cutting perpendicular to the lay of composite fibers, edge break-out can occur. This can be controlled by designing a backup structure in the tooling. 27.3 CONVENTIONAL TURNING
The turning of graphite composite is utilized to produce round surfaces that need to mate with either metal of graphite parts. The cutting speeds can be over 305 m/min (1000 ft/min) if the part can be held securely and PCD tool inserts are utilized.
Depth of cut will vary depending on the thickness of the part and the amount of material to be removed. 27.3.1 ADVANTAGES
Computer numerical controlled lathes (CNC) can be used to machine simple to very complex rotational parts. CNC machining produces accurate parts at a high production rate. 27.3.2 DISADVANTAGES
Delamination can also occur on a lathe (Fig. 27.1), therefore the part may require a finish cut moving from the largest diameter to the smaller diameter. Graphite chips are a serious problem. The spinning chuck creates a fan effect on the graphite particles. The exhaust system must be adequate to control the graphite chips. Also, the machine ways and the ball screws on the machine must have sealed protection to minimize wear. The computer control requires protection from the graphite chip particles. 27.4 MECHANICAL DRILLING OF COMPOSITE MATERIALS
Drilling holes in composites can cause failures that are different from those encountered when drilling metals. Delamination, fracture, break-out and separation are some of the most common failures. Delamination (surface and internal) is the major concern during drilling composite laminates as it reduces the structural integrity, results in poor assembly tolerance, adds a potential for long term performance deterioration and may occur at both the entrance and exit plane. Delamination can be overcome by finding optimal thrust force (minimum force above which delamination is initiated). Figure 27.2 shows push out delamination at exit because at a certain point loading exceeds the interlaminar bond strength and delamination occurs. Figure 27.3 shows peel-up delamination at entrance
598 Composite machining
Fig. 27.2 Drill bit showing push-out delamination at exit.
graphite reinforcement materials. Each of these materials requires individual attention in the selection of cutting tool parameters. The composite materials with metal backup panels require separate drills with different geometries. Cutting speeds and feedrates vary in each of the various combinations of materials. Secondary drilling or reaming operations are required to hold tight tolerances or smooth surface finishes on the holes. Table 27.1 shows Fig. 27.1 Machining direction for turning compos- the drilling results when using four styles of drills.' ite parts on a lathe. PCD tooling offers increased tool life, better hole quality, consistent hole size and higher because the drill first abraded the laminate machining rates. Drilling and countersinking and then pulled the abraded material away along the flute causing the material to spiral up before being machined completely. This type of delamination decreases as drilling proceeds since the thickness resisting the lamina bending becomes greater. Among the variables to be considered for tool selection include the thickness of material, diameter of hole, tolerance requirements, hole finish requirements and the composite materPeeling ial being drilled. Tungsten carbide, micrograin Action tungsten carbide and drill tool materials are used for drilling composite materials. Some commonly used composites are I glass-epoxy, glass-graphite-epoxy, graphiteepoxy, graphite-epoxy with aluminum backup and graphite-epoxy with titanium backup. Other materials include the aramids Fig. 27.3 Drill bit showing peel-up delaminationat (Kevlar@) with combinations of glass or entrance.
I
I
4 I
I I
Mechanical drilling of composite materials 599 Table 27.1 Summary of drill performance: mean hole quality measures as a function of point angle. Maximum recorded values of response parameters are shown in box brackets, [I (Reproduced from Ref 1 by permission of ASM Materials Week)
Criterion/drill
Dagger
8-Facet
4-Facet Master
NAS 907-1HSS
Exit breakout (Rank least = 1)
1
2
3
4
Panel damage, D,
1.96 (3.34)
2.37 (3.18)
2.75 (3.62)
3.63 (5.54)
1
2
3
4
Thrust force, N (1bf)
114 [166] (25.6 [37.4])
201 [378] (45.3 [85.2])
263 [428] (59.3 [96.3])
593 [969] (133.5 [218])
Torque, Nm (ft lbs)
1.29 [2.18] (0.95 [1.61])
1.15 [2.0] (0.85 [1.5])
0.7 [1.64] (0.50 [1.21])
1.53 [2.2] (1.13[1.61])
0.4 [1.6] (26 [641)
0.95 [2.2] (38 [88l)
1.6 [3.0] (64L1.221)
2.4 [4.12] (96 11651)
6.354 [6.379] (0.25016 [0.25115])
6.356 [6.369] (0.25022 [0.25075])
6.367 [6.395] (0.25067 [0.251751)
6.375 [6.397] (0.2510 [0.25185])
Microcrack density (Rank: lowest = 1)
Surface finish, R,, Pm (Pin.) Hole diameter, mm (in)
-
Hole out-of-roundness, (in)
Drill point angle, deg.
0.0061 [0.025] 0.003 [0.005] 0.0043 [0.018] 0.013 [0.03] (0.00024 [0.00101) (0.00012 [0.0002]) (0.00017 [0.0007]) (0.00051 [0.0012]) 30
with a combination tool provides better hole quality. Tool life is normally determined by the extent of delamination and fiber break out. For machining graphite composites with or without aluminum backing, PCD tooling is suggested with the same speeds and feeds used for machining graphite composites without any backing. For machining graphite composites with titanium backing, it is not recommended that the same drill be used for both the titanium and graphite sections. Initially a hole should be drilled up to the titanium layer with a hydraulic depth sensing device at high speeds and feed. A second drill with lower speed and feed for machining titanium should be used. Finally finish reaming operation and countersinking should be performed for assuring hole quality. A study carried out on carbon fiber-epoxy
24,118
140
135
(CFRP) and glass fiber-epoxy (GFRP) laminates using HSS and carbide tipped drills made the following observations. Both chisel edge and flank wear increased on the carbide drill with a higher ratio of wear between 200 and 400 holes (test sample 400 holes). The tool wear was greater in the CFRP laminates due to the abrasive nature of carbon fibers. Flank wear is more pronounced in GFRP when the feed was increased and the same effect is noted when speed is increased. The HSS drills lasted for ten holes in the graphite and twenty holes in the glass. 27.4.1 DFULL GEOMETRY
Drill point geometries influence the torque requirements. Lip relief and rake angles are determined by the application. The dagger drill is ideal to machine graphite composites
600 Composite machining as it eliminates breakout when exiting the workpiece. The dagger drill has 35" included point angle and a 121" chisel edge angle. Twist drills with flute configuration to control metal chips are also used. Fully fluted drills with PCD tips brazed on a solid carbide shaft provide the strength of carbide and hardness of diamond. Drill geometries are continuing to be experimented with to find ways to eliminate the problems associated with the hole making process in composites. Drill cutting parameters are: 0 0
0
feedrates range from 0.025 mm/rev (0.001 in/rev) to 0.063 (0.0025 in/rev); cutting speeds range from 30 surface m/min (100 surface ft/min) to 460 surface m/min (1500 surface ft/min); high cutting speeds can burn the matrix material and reduce bond strength between the composite material and the matrix material.
27.4.2 COOLANTS
A water soluble coolant forced through a cold air blast unit is recommended when machining most composite materials. However if the composite is hydrophilic in nature then a cold air blast unit in combination with dust or vacuum collection system should be used.
be processed. The grinding of polymer matrix composites (PMC) has a number problems. For example in the case of thermoplastic matrix, the surface of grinder becomes covered with melted thermoplastic. In the case of aramid fiber it is hard to get a clean cut surface because the grains cannot abrade the aramid fibers cleanly. Abrasive belts have been used on aramids with some success but dust collection has been a major problem. 27.6 MACHINING O F KEVLAR
Cutting, Trimming, Turning and Milling of Kevlar Because of its inherent toughness, Kevlar is difficult to cut, so sharp, heavy duty upholstery scissors will cut up to 170 g/m2 (5 oz/yd2)fabric of Kevlar. Woven roving and heavier fabrics can be cut using specially designed serrated scissors. An overview of cutting and trimming techniques and applications is shown in Table 27.2. For more information on cutting and machining of Kevlar refer to DuPont's Machining Handbook2. 27.7 ABRASIVE WATER JET MACHINING
Abrasive water jet (AWJ)is used for linear profile cutting, turning, milling and drilling operations in composite materials. 27.5 GRINDING COMPOSITE MATERIALS Conventional tool machining is affected by The grinding process has been used exten- fiber or particle reinforcements rather than the sively for finishing composite golf shafts and matrix material while AWJ machining is not. fishing rods. Five hundred parts per hour can To make a circular hole 6.35mm (0.25in) in be produced on centerless grinders. Silicon diameter in aramid 3.18 mm (0.125in) thick, it carbide wheels are used with an open grain to takes about the same time for both convenreduce wheel galling. Surface speeds between tional as well as AWJ machining. The cutting 1219 surface m/min (4000 surface ft/min) and process parameters for AWJ include water jet 1829 surface m/min (6500 surface ft/min) can pressure, velocity, abrasive grain size, abrasive be achieved. This equipment is specially material, standoff distance and jet impingedesigned for grinding and finishing compos- ment angle. and some additional parameters. ites. Grinding accuraces within 0.0127 mm Water jets without abrasive are also used for (0.0005 in) can be achieved with centerless cutting soft composites. Figure 27.4 shows the grinding. Both straight and tapered shafts can AWJ processes and machining parameters3.
0
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Abrasive water jet machining 601
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602 Composite machining
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Abrasive water jet machining 603 27.7.1 LINEAR CUTTING WITH AWJ
27.7.2 TURNING
Linear cutting is used to trim composite parts and to cut profile shapes on the inside of a part. The cut surface is normally smoother near the entrance surface then becomes wavy in the lower half of the cut toward the exit surface. In general, the composite material is sheared away by a high velocity abrasive grain. The width of cut (kerf) decreases as the feedrate increases and the waviness increases as the feedrate increases3. Table 27.3 shows some of the observations made by Hashish3.The maximum cutting traverse rate is primarily controlled by the matrix material. Table 27.4 shows results for some composites with different speeds.
In turning with AWJ, a workpiece is usually rotated while the jet is fed along all three axes. The material encountered by the jet is abraded away in the form of a very fine debris. Higher jet pressure produces a smoother surface with a higher material removal rate. Higher traverse rates combined with multiple passes are more efficient than deeper cuts with lower traverse rates. Surface finish is affected by unsteadiness in traverse rate or abrasive flow rate. The repeatability and accuracy of the AWJ turning process depends on control and steadiness. A 10% variation in rotational speed does not affect the surface waviness but a traverse rate variation over 4% will significantly affect the surface waviness. Some methods to improve surface finish are:
Table 27.3 Typical through-cutting traverse rates (in mm/s) with AWJs for different composites3 Material
Thickness (mm) 0.79
1.60
3.18
6.36
12.7
19.1
50.8
15 7.5 17 15 4.7
10 4 12 10 2.5
2.5 0.85 15 4.2 0.63
Organic matrix composites: Plastic and composites Carbon-carbon composites Epoxy-glass composites Graphite-epoxy composites Kevlar (steel reinforced)
53 42 105 74 42
38 32 95 63 25
29 22 76 52 17
21 13 42 40 8.5
Cutting conditions: p = 345 MPa, d, = 0.299 mm, d, = 0.762 mm, garnet mesh 80 Sic abrasives
Table 27.4 Surface waviness and corresponding cutting traverse rates (in mm/s) for some composite materials3 Material A1,OJSiC (20%) 6.35 mm Toughened zirconia (6.36 mm) Mg/B,C (15%) 6.36 I T U ~ Graphite+poxy composites (3.18mm thick) Graphite-epoxy composites (18.5 mm thick)
Rh4S surface waviness (pm) 1.9
2.5
-
0.15
-
0.6
3.8
5
6
8
-
0.29
-
-
-
-
0.2 3 4 0.85
Cutting conditions: p = 370 MPa, dn = 0.299 mm, d, 0.762 mm, garnet mesh 80
8 1.7
0.4 6 12 2.5
20 3.4
0.5 8 30 4.25
604 Composite machining 0
0 0
0
multipassing by traversing the jet without lateral feed; use of finer abrasive and increasing number of passes; to improve surface roughness, use softer abrasives like silica sand, copper slag etc; finishing with slurried abrasive yields improvement in surface roughness.
27.7.3 MILLING
The main objective of AWJ in milling is to produce a cavity with controlled depth. In this method, the jet material interaction is the depth determining factor. The production of kerf irregularity can be reduced by manipulating one of the factors, such as traverse rate, increasing the stand off distance or angling the jets. To mill square pockets the traverse speed can be varied rather than angling the water jet head. In this case the nozzle can be manipulated over the workpiece with an oscillatory drive using a motor and an eccentric. A uniform traverse rate and exposure time can produce a uniform depth cavity. A hard material pattern with the shape to be milled can be used to mask the target surface. This way the mask will allow jetting in the traverse zones where the traverse rate is uniform. Surface finish variations can be achieved by using different abrasive materials or grit sizes. Harder abrasives can be used for higher material removal rates and softer abrasives for finishing operations.
velocity decreases as the depth increases which can be attributed to the effect of return flow which reduces particle velocity and interferes with the impact process. Pressures of 3040 MPa are common for piercing glass. High pressures are necessary to pierce brittle or laminated composites. The higher pressures may cause the following problems: fracture due to shock loading of water; hydrocracking due to hole hydrodynamic pressurization; delamination due to loading. Holes larger than the piercing diameter of the AWJ are first pierced, then profile cut to the finished diameter being offset by the kerf amount. Hole shape variance depends on mixing tube length, target material, standoff distance, depth of hole and dwell time in the cut. Mixing tube length is important when drilling materials with high resistance. Increasing the mixing tube length improves the distribution of the abrasive with the water jet. This produces holes that are straighter and rounded. Advantages offered by AWJ are: 0 0 0 0 0 0
0 0
0
27.7.4 DRILLING
Hole drilling can be performed in any of the following ways depending on the diameter and accuracy of the holes: piercing is suitable for small diameter holes; kerf cutting is suitable for large diameter holes; milling is suitable for blind holes. Techniques of hole piercing vary for each composite material. Piercing glass, acrylic and polycarbonate show that the general geometrical features of pierced holes are similar. Particle
suitable for wide range of composites; can perform many operations like turning, drilling and milling; no thermal stresses; high as well as low material removal rates; no heavy clamping of workpieces required; omnidirectional machining; process can be automated; optimal range of parameters available to prevent delamination, loading and splintering; fine holes of 0.5 mm (0.012 in) can be drilled. Disadvantages:
0 0 0
0
dimensional accuracy is low; temperature rise in cutting region may be observed; limited data is available with respect to applications in metal and ceramic composites; not suitable for materials that are hydrophilic in nature.
Ultrasonic machining 605 Advantages:
27.8 LASER MACHINING OF COMPOSITES
Lasers are used in various industrial applications such as drilling, cutting, welding and heat treatment of metals, etc. In composites, polymer matrix materials are most suited for laser cutting. Laser cutting is a non-contact ablation process in which efficiency is determined by thermal properties of the workpiece material. Two types of laser have been used in industry: Nd-YAG solid state laser and CO, gas laser. The Nd-YAG laser operates in the near infrared (IR) region of the spectrum while CO, gas laser operates in the far infrared region. The Nd-YAG IR region wavelength is not absorbed by glass and many plastics while the CO, far IR region wavelength is. Applications of Nd-YAG solid state lasers extend from drilling fine holes in jet engines to welding implant devices for the medical industry. It has been determined that the NdYAG laser is very effective in cutting graphite-epoxy composite materials. The high power short pulses achieved with this laser vaporizes both the graphite and epoxy matrix before the epoxy resin can be overheated. The CO, gas laser applications extend from drilling holes in baby bottle nipples to welding automotive components in assembly lines. CO, lasers operate in either continuous wave or pulsed mode. Pulsed mode is preferred because of high powers obtained and cool down time. Aramid fiber reinforced plastic (AFRP) has been cut very effectively by the CO, lasers. The general characteristics of a laser cut zone in composite materials are shown in Fig. 27.5. The charred layer which includes a zone with fibers protruding from the matrix and as outer darkened zone in which the matrix has undergone some degradation4t5. Figure 27.6 shows the relationship between kerf width and cutting speed. For three dimensional (3D) machining two laser beams are directed through an optical assembly to intersect in the plane of work piece to cut shoulders and vee grooves.
0 0 0 0
0
superior quality edges due to high temperatures; vaporization of the material in cut zone; extremely localized action; sealing of the edge in the cut zone; pulsed CO, has been demonstrated as the best laser for processing Kevlar composites. Disadvantages:
0
0
0
beam divergence after its focal point; material thickness of about 9.5 mm (36 in) is the maximum thickness that can be cut with 1500 W; heat affected zone of varying dimensions.
27.9 ELECTRIC DISCHARGE MACHINING (EDMI
Advanced composites can be cut by EDM as there is no physical contact between the electrodes or workpiece and the tool. In order to EDM a composite, it should have an electrical resistivity of less than 1-3 ohm/m. Polymer matrix composite manufacturers can add a small amount of copper in the matrix of the product to allow shaping by EDM. EDM can be used with conductive silicides, borides, carbides, etc. The EDM process is more accurate than AWJ machining. Small holes of 0.25 mm (0.01 in) diameter can be drilled in SiC/TiB, composites. The EDM process is found to be slow for many production applications. 27.10 ULTRASONIC MACHINING
Ultrasonic machining (USM) incorporates a tool vibrating at 20 kHz and abrasive in a slurry to perform impact grinding of brittle materials. This technique is particularly useful for machining of ceramic matrix composites that are difficult to process by conventional methods. USM is a mechanical material removal process best suited for machining brittle materials like glass, ceramics, graphite and ceramic matrix composites. The process is limited to workpieces of size below 1OOmm
606 Composite machining
I
I
\ \
CHARRED LAYER
PROTRUDING FIBRES
\
I i
,i
\
,
'I
ICROSS SECTION
I
1
4 0 L-
beam exit side
Fig. 27.5 Schematic of FRP laser cut. (Reproduced by permission of Marcel Dekker Ltd.) W,: kerf width at the beam entry side; W,: kerf width at the
02 -
-
Fig. 27.6 Kerf width as a function of cutting speed for (0/90), laminates. (Reproduced by permission of Marcel Dekker Ltd.)
'.'a
'
m
u
'
)
Bo
80
Cutting speed (mm/s)
Irn
120
1
Ultrasonic machining 607 (3.94 in) because of the limitation on the size of the tool. Some of the variables that influence USM for close tolerances are as follows:
Abrasive type and size Abrasives contained in the slurry do the actual machining so they must be selected on the basis of the workpiece material and the surface quality needed. As in the case of AWJ, larger abrasive grains give higher material removal rates and rougher surfaces. The grain diameter cannot be larger than amplitude of the sonotrode as this would inhibit the injection of the grains to the machining gap. Common types of abrasive used are A1,0, oxide, Sic, BC and diamond. Table 27.5 shows recommended abrasive for various materials. The grain diameter affects surface roughness, overcut and machining rates. When high removal rates are necessary with no high surface quality required, 180-280 mesh abrasive do the job. For finer surface finish 320-600 mesh abrasive is recommended. Table 27.6 shows surface roughnesses for different workpiece materials.
Sonotrode (tool) material Tools with diamond tips have good material removal characteristics and very low wear but are difficult to machine. Table 27.7 shows accuracy results of using a non-rotating steel sonotrode. Ultrasonic vibrations The ideal condition would be the amplitude of ultrasonic vibration to be equal to the grain mean diameter. If the amplitude is too small the abrasive cannot enter the machining gap, if too large it causes the grains to be incorrectly projected. A mixture of both the types of abrasive may be used and a suitable amplitude selected to determine which size grain enters the machining gap. Surface area This factor influences removal rates and tool wear. With a small diameter, higher feed rate is obtained but also higher tool wear is noticed. This can be overcome by using a diamond tool or with a closed loop force sensitive
Table 27.5 Recommended abrasive for various materials6 Material
Recommended abrasive
Graphite Zirconia
Silicon carbide Silicon carbide or boron carbide Silicon carbide Boron carbide
Ceramic matric composites Metal matrix composites
Table 27.6 Surface roughness for various materials6 Workpiece material
Graphite Zirconia Ceramic matrix composites Metal matrix composites
Surface roughness Ra @ m) 1-2 0.75 0.70 0.90
608 Composite machining servo system maintaining accurate machining on prepreg materials like glass fiber, carbon pressures. Table 27.8 shows typical ultrasonic fiber and Kevlar with reduced fiber damage. Advantages: machining rates for a variety of materials6. USM is used in applications like drilling aerospace cooling holes in ceramic matrix 0 conductive and nonconductive materials can be machined; composite turbine blades, slotting, irregular 0 material hardness is not so important; configurations in ceramics and composites, machining of phased array radar components, 0 there are no chemical or electrical alterations in the workpiece; cutting tool inserts, superconductors, wire 0 3D and complex shapes can be machined draw dies and extrusion dies. A CNC USM can easily and quickly; cut through 6mm (0.24in) thick composite 0 no heat affected zone. layers and produce a controlled depth up to 50mm (1.97in). The latter is important, as Disadvantages: many composites have backing sheets that should not be damaged. The ultrasonic action 0 amplitude of ultrasonic vibrations are very important for proper machining; reduces the amount of force required to sever 0 limited sizes can be machined. the hard materials. This results in a better cut
Table 27.7 Accuracy results with a non-rotating steel sonotrode6
Material
Inlet diameter (mm)
Outlet diameter (mm)
Taper (Yo)
Roundness (mm)
Graphite
10.23-10.25 10.26-10.29
10.07-10.10 10.02-10.05
3.00 2.70
0.03" 0.03b
Metal matrix composite
10.20-10.24 10.09-10.12
8.87-9.92 8.85-9.90
9.00 6.60
0.04b 0.05b
Ceramic matrix composite
10.11-10.15 5.04
10.00 4.99
3.50 1.25
0.04b
5.05
4.85
5.50
-C
Zirconia
-c
Tool 1: Exponential,Diameter = 10 mm Tool 2: Exponential,Tube D = 10 mm, ID = 7 mm ' Tool 3: Exponential,Diameter = 5 mm a
Table 27.8 Typical ultrasonic machining rates for a variety materials7
Drilling diameter = 5 mm
Drilling diameter = 10 mm
Material
Time (min)
Removal rate (mm3/min)
Time (rnin)
Graphite Ceramic matrix composite Metal matrix composite Zirconia
1 3.5 10 210
164 39 7.6 0.65
1.25 5.6 14 90
Removal rate (mm3/min) 224 50 9.3 3.1
References 609 REFERENCES 1. Mehat, M., Reinhart, T.J. and Soni, A-H., Effect of fastener hole drilling anomalies on structural integrity of PMR-l5/GR composite laminates, Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Ill, 1-5 Nov. 1992. 2. Kevlar Cutting and Machining Handbook, E.I. Du Pont de Nemours and Co. 3. Hashish, M. State of the art of abrasive waterjet machining operations for composites. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. 4. Di Ilio, A., Tagliaferri, V. and Veniali, F. Machining parameters and cut quality in laser cutting of aramid fibre reinforced plastics. Materials and Manufacturing Processes, 1990,5(4), 591-608. 5. Lemma, S. and Sheehan, B. Laser Machining of Composite Materials. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. 6. Gilmore, R. Ultrasonic machining of composite materials, Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Ill. 1-5 November, 1992.
FURTHER READING Bhattacharyya, D., Allen, M.N. and Mander, S.J. Cryogenic Machining of Kevlar Composites. Materials and Manufacturing Processes, 1993,8(6), 631,651
Bhatnagar, N., Naik, N.K. and Ramakrishnan, N. Experimental investigations of drilling on CFRP composites. Materials & Manufacturing Processes, 1990, 5(4), 591-608 Geskin, E.S., Tisminetski, L., Verbitsky, D., Ossikou,V., Scotton, T. and Schmitt, T. Investigation of waterjet machining of composites. Proc. Machining of Composite Materials Symy., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. Hochegn, H., Puw, H.Y. and Yao, K.C. Experimental aspects of drilling of some fiber-reinforced plastics. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. Krishnamurthy, R., Santhanakrishnan, G. and Malhotra, S.K. Machining of polymeric composites. Proc. Machining of Composite Materials Symp., ASM Materials Week, Chicago, Illinois, 1-5 November 1992. Lubin, G., ed., Handbook of Composites, 1982, New York: Van Nostrand Reinholt. Ramulu, M., Faridnia, M., Gargini, J. L. and Jorgensen, J. E. Machining of graphite/epoxy composite materials with polycrystalline diamond (PCD) tools. Trans. ASME, J. Engng Mater. and Tech., 1991,113, October . Zaring, K., Erichsen, G. and Burnham, C. Procedure optimization and hardware improvements in abrasive waterjet cutting systems. PYOC. Machining of Comr?osite Materials Svmp., ASM Materials" Week, Chicago, Ill., 1-5 "November 1992.
MECHANICAL FASTENING AND ADHESIVE BONDING
28
D. W. Oplinger
28.1 INTRODUCTION
It would be difficult to conceive of a structure that did not involve some type of joint. Joints often occur at a transition between a major composite part, where most of the structural performance is generated, and a metal feature, which is introduced to allow for very high localized bearing contact for which the composite has inadequate strength or durability. In aircraft such a situation is represented by articulated fittings on control surfaces as well as on wing and tail components which require the ability to pivot the element during various stages of operation. Tubular elements such as power shafting often use metal end fittings for connections to power sources or for articulation at points where changes in direction are needed. In addition, assembly of the structure from its constituent parts will involve either bonded or mechanically fastened joints or both. Joints represent one of the greatest challenges in the design of structures in general and in composite structures in particular. The reason for this is that joints entail interruptions of the geometry of the structure and often material discontinuities, which almost always produce local highly stressed areas, except for certain idealized types of adhesive joint such as scarf joints between similar materials. Stress concentrations in mechanically fastened joints
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
are particularly severe because the load transfer between elements of the joint have to take place over a fraction of the available area. For mechanically fastened joints in metal structures, local yielding, which has the effect of eliminating stress peaks as the load increases, can usually be depended on; such joints can be designed to some extent by the 'P over A' approach, i.e. by assuming that the load is evenly distributed over load bearing sections so that the total load (the 'I") divided by the available area (the 'A') represents the stress that controls the strength of the joint. In organic matrix composites, such a stress reduction effect is realized only to a minor extent, and stress peaks predicted to occur by elastic stress analysis have to be accounted for, especially for one-time monotonic loading. 28.2 MECHANICALLY FASTENED JOINTS COMPARED WITH ADHESIVE JOINTS
In principle, adhesive joints are structurally more efficient than mechanically fastened joints because they provide better opportunities for eliminating stress concentrations; for example, advantage can be taken of ductile response of the adhesive to reduce stress peaks. Mechanically fastened joints tend to use the available material inefficiently and are characterized by sizeable regions where the material near the fastener is nearly unloaded, which must be compensated for by regions where high stresses occur to achieve a particular
Mechanically fastened joints 611 required average load capacity. As mentioned above, certain types of adhesive joint, namely scarf joints between components of similar stiffness, can achieve a nearly uniform stress state throughout the region of the joint. In many cases, however, mechanically fastened joints cannot be avoided because of requirements for disassembly of the joint, for replacement of damaged structure, or to achieve access to underlying structure. In addition, adhesive joints tend to lack structural redundancy, and are highly sensitive to manufacturing deficiencies, including poor bonding technique, poor fit of mating parts and sensitivity of the adhesive to temperature and environmental effects such as moisture. Assurance of bond quality has been a continuing problem in adhesive joints, primarily because, while ultrasonic and X-ray inspection may reveal gaps in the bond, there is no present technique which can guarantee that a bond which appears to be intact by ultrasonic or Xray inspection does not lack load transfer capability, because of such factors as poor surface preparation. Surface preparation and bonding techniques have been well developed, but the possibility that lack of attention to detail in the bonding operation may lead to such deficiencies needs constant alertness on the part of those responsible for the bonding. Thus mechanical fastening tends to be preferred over bonded construction in highly critical and safety rated applications such as primary aircraft structural components, especially in large commercial transports, since assurance of the required level of structural integrity is easier to guarantee in mechanically fastened assemblies. Bonded construction tends to be more prevalent in smaller aircraft however and for non-aircraft applications as well as in non-flight critical aircraft components.
composites first came into use. It was found early in this period that the behavior of composites in bolted joints differs considerably from that which occurs with metals, primarily because stress concentrations are much more of a factor in joint behavior of composite structures, and stress analysis to quantify the level of various stress peaks is more important. It was fortunate that significant computing power became available in this period to keep up with the need for the intimate details of stress conditions around mechanical fasteners. The current approach to the design of mechanically fastened joints in composite structures evolved mainly out of a number of DoD, NASA and associated university programs aimed at providing a methodology which could be applied routinely to aircraft and other applications. Numerous stress analysis approaches to the mechanics of fastened joints have been conducted over the years since the introduction of 'advanced' composites in the mid-1960s. These have included: the work of Waszczak and Cruse' based on the boundary integral method; the use of two-dimensional complex variable elasticity solutions which treated the problem of variable contact around the fa~tenel-2-~; as well as recent work reported by Madenci and Illeri7; and a number of finite element approaches, especially the work of Crews and Naik8 which featured an inverse method for dealing with the contact problem. Hart-Smith9 developed an analytic approach based on the use of available solutions for isotropic plates with bolt-loaded holes as well as unloaded holes in plates under tension or compression which came out of classical efforts such as those reported in Petersonlo.The latter provided simple functional descriptions of the effect of joint geometry on peak stresses which, with various empirically derived correction factors introduced by Hart-Smith9, provided valuable insight into a number of trends in joint 28.3 MECHANICALLY FASTENED JOINTS behavior. In addition to the analytic efforts, several Mechanically fastened joints for composite structures have been under study since the fairly extensive programs aimed at the develmid 1960s when high modulus, high strength opment of design approaches for structural
612 Mechanical fastening and adhesive bonding systems have been supported by DoD and NASA4r6.Numerous papers have been presented Over the years in a series Of DoD/NASA/FAA 'Onferences On Fibrous Composites in Structural De~ign~,'~-'~. Many of the design principles which have been devel'ped to take into account the characteristics of bolted composite structures have been described4,6, 13. It is not possible within the scope Of this discussion to describe all the details and Processes that are necessary for achieving the design of specific joints. The objective is rather to give the reader Some insight into the factors that control the behavior of mechanifastened joints in structures' The behavior of mechanically fastened types Of information: joints is governed by (l)the features that the behavior Of the joint around individual fasteners and; (2) the behavior of multiple arrays of fasteners. The behavior of individual fasteners can be considered in terms of a generic rectangular element surrounding each fastener (Fig. 28.1), whose length and width are represented as ratios wi& respect to the fastener diameter. The effects of the geometry of this element together with effects of the reinforcement arrangement used in the laminate for the element determine the structural conditions under which the element will fail. Once the characteristics of the rectangular element are selected, its deformation characteristics can be combined with those of other rectangular elements making up the joint to obtain the performance of the joint as a whole. This discussion is organized in terms of these two aspects of joint design. In addition to a preliminary discussion of general features of mechanically fastened joints, the discussion which follows also considers: (1) single fastener joints, including effects of joint geometry together with those of composite material behavior; (2) multi-fastener joints; (3) fastener effects, and (4) a discussion of test methods which provide empirically-based data needed for completing the joint design.
28.3.1 KEY FEATURES OF MECHANICALLY FASTENED JOINTS
~~
Figure 28.1 represents a generic single fastener joint, while Fig. 28.2 depicts a multiple fastener configuration. Many of the most important features of joints are illustrated in the single fastener case s h o r n in Fig. 28.1. Key dimensions are D, the fastener diameter, t, the thickness of the joint structural elements, e, the edge distance (distance from the fastener center to the edge of the upper plate) and W, the width of the part of the upper plate to the left of the fastener. Similar featups apply to the lower plate element. Note that the Same load, p, is passed successively through various sections of the joint, including the bearing section in front of the fastener of area Dt which is in compression, the b o shearout sections of total area 2et which are loaded in primarily in shear, the net section, - D)t which is in tension, the gross section wt which is also in tension. The average stresses associated with these sections are:
(w
ab= P / D t ; average bearing stress, average shearout stress, a,, = P/2et; average net section stress, aN= P/(W - D)t; average gross section stress, aG= P / Wt. (28.1) LOAD PATH
P t
I
'
goss section
\
bearid
sectiqn
net section (W-O)t
wt
Dt
I I I
I I
,/-7-. c
I
I
!
:
I
Fig. 28.1 Single fastener joint.
I
J
section
Mechanically fastened joints 613 Axit!
tit h
w
Lateral Pitch
eff
*
Lateral Pitch
I
Fig. 28.2 Multi-fastener joint.
From the standpoint of the designer, the gross section strength is of primary interest since the objective of good design is to stress the gross section to its highest level. Structural performance of the joint can be rated in terms of joint efficiency, which refers to the ratio of average gross section stress at failure of the joint to the strength of the laminate in the gross section, essentially the strength achieved in coupon tests for tensile or compressive loading of unnotched specimens. For organic matrix composites, single fastener joints achieve joint efficiencies of less than 50%, while for multifastener joints the maximum achievable efficiency is of the order of 60%. In contrast, metallic joints can reach efficiencies of close to 80% because of the opportunity for taking advantage of local yielding around points of high stress, though even in metallic joints, design for avoidance of crack initiation in cases where long life under cyclic loading is required may force the joint efficiency to be lower than for single cycle loading.
-8-1
G
Fig. 28.3 Peak stresses around fastener.
gross section:
KGt = O,&JO~
(28.2)
or
qt= q J m x / o G
where , ) a is the maximum axial stress on the net section. Predictions of the peak stresses in the joint can be made using continuum elasticity analysis1", 7, 12, photoelastic measurements'O or finite element methodsE. 28.3.2 EFFECTS OF JOINT GEOMETRY The behavior of K$ as a function of D / W The peak value of axial stress on the net sec- for isotropic joint elements is given in Fig. 28.4 tion on,),,, (Fig. 28.3) is one of the key which was obtained from photoelastic meacontrolling parameters on joint performance. surements of KNnt against D / W (note that Kgt = It is convenient to express this as a stress con- KNnt/(l - D / W ) ) given by Petersenlo. Similar centration factor, i.e. a ratio with respect to curves for orthotropic plates were obtained by either the mean stress in the net section or the elasticity analysis2,3, 12.
614 Mechanical fastening and adhesive bonding There is obviously a competition for space between the bearing section and the net section. Increasing the fastener diameter, D, to lower abhas to decrease (W - D)t, the net section, thereby increasing uN, and vice versa. Furthermore, if D / W is much less than 1, the case is similar to that of a fastener in an infinitely wide plate; in such a case all the peak stresses become a constant multiplier of the bearing stress ab, which becomes large for D/W<
Kgt
0.89W/D
+ 1.8/(1 - D/W)
'T
The behavior of this approximation emphasizes the tradeoff between the effect of the bearing section and the net section on the peak net section stress. A similar functional fit to the curve of KNt (the stress concentration based on a,) against D / W was given for isotropic plates by Hart-Smith9, based on photoelasticity results reported in the literature; when stated in terms of Kgt, the Hart-Smith approximation leads to
K;t (bolt loaded hole) = W/D
- 1.5/ (1+ D / W )
+ (1 - D/W)*
Functional Fit, 0.89%+ 1.8oN
4 --
01 0.1s
elastic measurements)
I
028
(28.4)
Kgt (open hole) = 2/(1- D / W )
(28.3)
3 --
D/W)
He also gave an expression for KNt for open, unloaded holes (i.e. with no load transfer between the fastener and plate) in isotropic plates; when the stress concentration is expressed as a ratio with respect to ac rather than uN the Hart-Smith open hole expression reduces to:
\ Y
+ 2/(1-
0.53
0.4
DIW
---31.-
Fig. 28.4 Predicted net section stress concentration against D /W.
0.65
(28.5)
Mechanically fastened joints 615 This is of interest in later discussions of com- delamination and fiber-separation around bined bearing-bypass loads on bolted joints. regions of peak stress. The method is applied As a means of predicting joint strength, we can by averaging the stresses predicted by twoequate the peak stress CTJ,,,~~ to F'", the dimensional elasticity analysis over a selected strength of the orthotropic laminate used in length a,, along a path lying normal to the load the joint. If the joint material acts purely elasti- direction, in the vicinity of the peak stress. The cally, then the elasticity prediction of K$ gives failure load is obtained by equating the averF; which denotes the value of C T when ~ net sec- aged stress to the laminate strength P, and u, tion failure occurs, as F," = F'"/K,"'. The is determined by an empirical fit of the meaminimum in the KGntagainst D/W curve of sured strength data. In an alternative version of Fig. 28.4 will result in a maximum in the rela- the Whitney-Nuismer analysis, F'" is equated tionship between F," and DIW at D / W = 0.5. to the stress at a point spaced an empirically fitHowever, joint strength is usually found to ted distance do from the location of the stress be considerably greater than that based on peak. The latter approach has frequently been purely elastic behavior of the composite to fail- used for joint strength predi~tion'~, although ure, and it is appropriate to use the method the averaging and point stress versions of the developed by Whitney and NuismerI4 to cor- Whitney-Nuismer analysis are equivalent. rect strength predictions based on linear elastic Figure 28.5 illustrates the prediction of F; response to failure. The Whitney-Nuismer for net section tension failure as a function of approach was developed for strength predic- D / W . The curves here were obtained from tion of composite plates in the presence of stresses given by elasticity analysis of 0/90° circular holes, crack-like slits and other types of glass epoxy joint elementsI2,by averaging the notches, and is meant to take into account the predicted axial stress distribution uxin the net softening effect of local damage such as section (usingthe definition of x and y indicated
D = 0.63 cm (0.25 ") Laminate Tensile Strength 110 ksi (0.76 GPa) 45 ,-0.31
Compressive Strength 100 ksi ( 0.69 GPa)
GPa
40 0-
vi
35 30
c
C
50 $ g
25 20
15
v)
10
L
a 5 0 0.13
0.25
0.33
DIW Fig. 28.5 Maximum gross section stress against D/ W.
0.5
0.63
616 Mechanical fasfening and adhesive bonding by the coordinate axes centred at the hole in Fig. 28.3 from the edge of the hole to a point a distance u, from the hole, using the a, values listed in Fig. 28.5. The lowest curve of Fig. 28.5 which corresponds to a, = 0 represents purely elastic behavior of the joint material. It is clear that increasing the value of a, causes an upward shift of the predicted strength curve. Experimental joint strength datal2 gave an a, value of about 3.8 * 1.3 mm (0.15 & 0.05 in) for quasi-isotropic carbon epoxy laminates, for net section failure. Nuismer and W h i t n e ~ ' ~ suggested that a, and the corresponding d, are relatively constant for composite materials in general, although this is not always true. However, for a given material and stacking sequence, a, is independent of hole size; this M has the effect that for large fastener diameters, the averaging effect is less important, and Fig. 28.6 Experimental results on strength against D/W2. strength values tend toward those predicted by the elasticity analysis. All the net section failure curves of Fig. 28.5 'bearing failure', corresponding to compressive have the peak in the vicinity of D/W = 0.5 failure in front of the pin. Except for joints with which was predicted by the linear elastic short edge distance or multi-fastener joints response analysis. This suggests that the joint where high bypass loads are present, the peak strength is a maximum for bolt diameters near compressive stress in front of the fasteneP is half of the plate width. For multi-fastener found to be about 1.3 times ab,and uGfor bearjoints such as the one shown in Fig. 28.2, the ing failure can be estimated by setting abequal length corresponding to W for the single fas- to F" (compressive strength of the laminate) tener joint is replaced by WeE, the spacing of from which uG)max = F'" D/1.3W, a linear functhe dashed lines, which are lines of lateral tion of D / W. Although this result implies that symmetry; note that We* is also equal to the the laminate compressive strength is the same lateral s cing of the fasteners (the lateral as the bearing strength as is approximately true pitch), so t t the peak strength value occurs in some cases, the actual value of Fb" = ub)mm at a lateral pi ch of about twice the fastener depends on the stacking sequence of the lamidiameter. The experimental data of Fig. 28.612 nate as well as a number of details of the joint confirms the occurrence of a strength maxi- such as the type of fastener and whether the mum for D / W = 0.5 in single fastener joints. joint is in single or double shear (Section For multi-row joints (see Section 28.3.4) in 28.3.5); thus Fbu has to be empirically deterwhich bypass load is present, the net section mined for a given fastener, joint configuration strength peak occurs for smaller values of and joint material selection by appropriate D / W. High bypass loads give results similar bearing strength tests (Section 28.3.6). Hartto those for plates with unloaded holes, in Smith9J3introduced a method of comparing which case large values of W / D correspond- bearing and net section tension failure similar ing to large fastener pitch are desirable. to that given in Fig. 28.5 to show the trade off In addition to the solid curves in Fig. 28.5 for in failure modes for various values of D/W, net section failure is the dashed line designated and recommended selecting D/ W for the
-
\
Mechanically fastened joints 617 design of the joint as the point at the intersection of the bearing and tension failure envelopes. However, it was also pointed out in Nelson, Bunin and Hart-Smith13 that in multirow joints in which the bearing load is distributed over several fasteners in a given column, the bearing stress is smaller than for the single fastener case and bearing failure may not occur in practical joint designs. The value of the edge distance, e, is another critical parameter for bolted joints. While it has become customary to present joint strength as a function of e / D , in fact9,12it is more a function of e/ W if e / D is much greater than 1, and plotting data on the effect of edge distance as a function of e / D is somewhat misleading even though it persists as accepted practice. Both stress analysis9,l2 and experimental results (Fig. 28.7) indicate that full joint
strength will not be achieved unless e is at least as great at W. This is shown schematically in Fig. 28.8, which indicates that the minimum e for full joint strength over a wide range of fastener diameters remains equal to W. It is true that there is a tendency for bending failures to occur at the unloaded end of the joint (Fig. 28.9) for edge distances as small as 1-2 times D, which were shown in Op1ingerl2, in fact, to be a function of e / D; however, this type of failure is eliminated for e / D greater than about 2.
Fig. 28.8 Minimum edge distance is equal to W (independent of D).
y /
I
high bending
stress 0
Y
b
I
I
I
1
I 4
.W
Fig. 28.7 Effect of edge distance on joint strength W / D = 212.
Fig. 28.9 Tendency for bending failure with small e/ D.
618 Mechanical fastening and adhesive bonding 28.3.3 EFFECTS OF JOINT MATERIAL
Typical failure modes for mechanically fastened joints are shown in Fig. 28.10. Figure 28.11 indicates that shear-related failures may include what are usually 'shearout' failures such as the upper figure for 0,&5 reinforcements, which tend to lie along straight lines tangent to the bolt hole and to run to the end of the joint, as opposed to those for 0/90° laminates which originate at about 45" points around the bolt hole and run along curving lines to the joint end. Hart-Smith9 shows examples of 'shearout' type failures while the second type are demonstrated by experimental results for 0/90" glass epoxies12.Since the 90" points around the bolt hole are traction free, shear stresses cannot exist there, and 'shearout' failures are probably associated with fiber separation failures resulting from to a deficiency of 90" fibers in 0,&5 laminate, while failures originating at the 45" points in 0/90° laminates appear to be true shear failures. Chang and Hun@ provided an analytical approach which modelled progressive damage in 0/90° laminates that predicted the types of shear failures that are depicted in Fig. 28.11.
Fig. 28.11 Shear failures in bolted joints.
In addition to modelling shear failures, Chang and Hung16 also reported results related to bearing failure which show that these failures result from high out-of-plane shear stresses; the associated failure surfaces are oriented along planes at 45" to the x-y plane. In addition to clarifying the bearing failure modes, these results verified the importance of clamping pressure as a factor in achieving maximum bearing strengths. This is discussed further in Section 28.3.5. It should be mentioned that the term 'bearing' strength, which is usually associated with the maximum achieved value of ab,does not necessarily imply that the joint fails by bearing failure, since, depending on the joint geometry
l€NSlON FAIUIIL
Fig. 28.10 Typical failure modes in composite joints13.
BO11 PULLING lHlOUGn U M I W T E
Mechanicallyfastened joints 619 and reinforcement, failure may take the form of shearout or net tension rather than bearing failure; bearing strength only identifies the value of abthat occurs when the joint fails. The value of abfor the bearing failure mode can be looked on as a laminate material parameter, however, and can be obtained from appropriate single fastener strength tests, to be discussed in Section 28.3.6, in which e/ W (usually specified in terms of e / D ) and W / D are large enough to guarantee a bearing failure mode. The effect of the reinforcement configuration on the maximum achieved ab is an important factor in joint design. Note that for design purposes, the normal reinforcement arrangement includes only plies in the 0", 45" and 90" directions with respect to the load axis, implying that other orientations are not encountered; in addition, the 45" reinforcement is arranged symmetrically as double plies of k45" reinforcement. The notation commonly used to describe the stacking sequence is condensed to reflect percentages of plies in the three customary directions, 25/50/25 representing 25% 0", 50% 45", and 25% 90", or quasi-isotropic, reinforcement, for example. The effects of reinforcement percentages on maximum achieved bearing stress are shown in Fig. 28.12, a three-parameter 'carpet plot' which is used to show the effect of percentage of reinforcement in the 0", 90" and 4 5 " directions on maximum achieved bearing stress for a fixed joint geometry. The contours of constant bearing strength indicate that maximum joint strength is achieved within a broad plateau lying between 3040% 0" and 1040% +45" reinforcement. To avoid low strength against shearout failure, it is generally agreed that a minimum number of 90" plies, at least lo%, should be included in the laminate. Laminates reinforced in fewer than three directions, i.e. 0/90° and k45" reinforcement, should be avoided if possible, since they respond ductiley with excessively low yield strengths. These recommendations may be difficult to follow if considerations other than maximum joint structural performance govern
PEICfNTAGf W-DEGIEE PLIES
BEARING STRESS CONTOURS 6 5 1 )
WLT DIUETER = 0.25 IN. (10RQUED WLW
O
M
Z
O
3
D
~
S
O
~
M
~
W
RKfNlAGE f4S-DEGltE PLIES
Fig. 28.12 'Carpet' plot of maximum bearing stress vs. percentage of reinforcement in key directions4.
the reinforcement arrangement in the vicinity of the joint. 28.3.4
Single fastener joints cannot generally achieve anything close to the strength of the laminate being loaded and are not usually encountered in structural joints. Single fastener coupons do, however, serve as building blocks for design of multi-fastener joints (Fig. 28.2) since stress analysis and strength test results on single fastener geometries provide data which translate directly to multi-fastener arrays in structures. There are two principal features that affect multi-fastener arrays; that of bearing-bypass load is illustrated in Fig. 28.13. The bypass load, Pb , can be considered as a load added at the d o a d e d (right) end of the upper element in the single fastener joint in Fig. 28.1. In the case of a multi-fastener array (Figs. 28.2, 28.14), Pbpfor a given fastener is equivalent to the sum of the loads developed by the fasteners lying to the right of (i.e. those lying in the direction of the load from) the one under consideration. For each fastener, the load P, in
l
W
620 Mechanical fastening and adhesive bonding On the other hand, the net section load on the
pN Pf +pbp
Fig. 28.13 Definition of bypass load.
Fig. 28.13 (see also Fig. 28.14) is the load passed through the net section associated with that fastener and is the sum of P , the fastener load, and Pb ;P, is also equivalent to the total load at the feft end of the joint in Fig. 28.13. The situation in Fig. 28.14(b) shows the buildup of P, along a joint containing five fasteners for the hypothetical case of equal P, for each fastener, while Fig. 28.14(c) shows the distribution of Pb /PNfor that case. It is important to realize that the net section stresses build up along with P,. If the joint design is such as to distribute the fastener loads P, evenly as in Fig. 28.14, the load on each fastener is then the total load divided by the number of fasteners, a fraction of 1/5 in Fig. 28.14, so that the bearing stress at each fastener is much smaller than for a single fastener joint.
Fig. 28.14 Bypass against bearing load.
last fastener is the sum of all fastener loads and is therefore the same as in a single fastener joint carrying the same total load. The benefit of multi-rowed fastener configurations is primarily because the stress concentration factor associated with pure bypass load (eqn (28.4))is smaller than that associated with pure bearing load (eqns (28.7), (28.3)),and the load ratio for the most highly loaded net section (the left most in Fig. 28.14) tends toward the case of pure bypass loading for large numbers of fasteners. Stress analyses for combined bearing-bypass loading have been discussed7,*, 12, 17, la. Figure 28.1512 gives the elasticity predictions of KEt for various ratios of Pbpto P,, indicating the KEt is a nonlinear function of P,/P . (The results in Fig. 28.15 are for a perbF fect fit fastener and for a joint under tension.) This nonlinearity is caused by the variation of angle of contact between the fastener and the plate with the load ratio; Fig. 28.16 illustrates the difference in the contact region for the extreme cases of pure bearing and pure bypass load. In the latter case the plate stretches in opposite directions along a line parallel to the load direction and contact splits into two regions centered about +90" and -90" with respect to the load direction, while for pure bearing load, a single contact region occurs ranging from about -100" to +loo" about the load axis for perfect-fit fasteners. (For clearance fits in the case of pure bearing load the
Fig. 28.15 Net section SCF against bearing/bypass load ratio.
Mechanically fastened joints 621
[AI
contact
WWS
-
ce)
-
-
BEARING
Fig. 28.16 Contact angles for pure bearing load and pure bypass load.
contact region varies with load and is smaller than that of exact-fit fasteners. Crews and Naik8 treated the case of 1.2% clearance, i.e. a hole diameter 1.2% greater than the fastener diameter, for which contact between the fastener and hole occurs from about -60" to +60° for typical loads.) For intermediate bearing/bypass ratios, the contact region is a mix of the two situations. Oplinger12gave the variation of the radial pressure distributions for exact fit fasteners, for joints in tension, which showed how the contact region is modified as the bearing/bypass ratio varies, while Naik and Crewss treated cases of both exact and clearance fits for both tension and compression loaded joints; in addition, Madenci7 gave comparable results for cases of shear loaded joints. A number of efforts have treated the contact problem for pure bearing by assuming a radial pressure distribution which varies around the hole as the cosine of the angle with respect to the load axis; in addition, for combined bearing and bypass loading (following industry practice) Hart-Smith9 and others have, as a matter of expediency, superposed the peak stresses predicted for pure bearing and pure bypass (i.e. the values of Kt: predicted by eqns (28.4) and
(28.5)). The results, given in Oplinger12,indicated that this will result in inaccurate predictions, not only for the general bearingbypass situation but also for the case of pure bearing load with small edge distances (Fig. 28.9). Currently available analytical and finite element tools are sophisticated enough to treat the contact problem routinely, and the 'cos 8' radial pressure distribution should be avoided, although the superposition method gives some useful insight into the situation if analytical tools for dealing with the contact problem are not available. Crews and Naiks obtained results which showed that the hoop stresses around the bolt hole are predicted with reasonable accuracy by the superposition method, so that with judicious use of N~ismer-Whitney'~ correction factors, net section failure stresses can be reasonably well predicted by superposition; bearing failures cannot. Naik and CrewsIs described test methods for joint strength under combined bearingbypass loading, with compressive as well as tensile bypass loads. Typical results are given in Fig. 28.17 for a 1.2% clearance fit fastener in a quasi-isotropic carbon epoxy laminate. Failure modes here are designated 'TRB' ('tension reacted bearing', i.e. bearing failure with tension bypass loading), 'TRC' (bearing 'failure with compression bypass loading), 'NT' (net section tension), 'NC' (net section compression) and 'OSC' or 'offset compression' which refers to the failure mode illustrated in Fig. 28.18 for compression bypass load. Although the load distribution for the fivefastener joint shown in Fig. 28.12 is represented as having the same P, for each fastener, this condition cannot be achieved for joints in general. As shown in Fig. 28.1913, there is usually considerable variation of fastener loads along the joint. In a two-fastener joint such as that shown in Fig. 28.20, the upper and lower plate elements ('U' and 'L') have to stretch equally under load if there are no fastener deflection effects (no fastener tilting or bending).
622 Mechanical fastening and adhesive bonding Bearing stress, Sb, Wa
Bypass stress, S
nP'
Wa
Fig. 28.17 Laminate strengths under combined bearing/bypass loading17.
and EGL
= gGL/EL,
the elongations are:
6,= kGu and 6 , = ZE,,. The condition that 6, = 6, requires that:
I(P - P,,)/E,t,W
= lPf,/E,f,W.
Since Pp = P - P,, this is equivalent to:
Pf, = (E,f,/E,t,)Pf,, If E,>>E, or tL>>t,, the resulting P,, will be small compared with P,, so that the second fastener will be nearly unloaded. In such a case With the average gross section stresses and the joint will be equivalent to a single fastener strains for the two plate elements in Fig. 28.20 joint containing an unloaded hole at the locagive by: tion of the second fastener. On the other hand, OGU= ( P - P, )/tuW; the last equation showsthat if the stiffnesses of the upper and lower plates are equal, the two fasteners will be loaded equally. In general, fastener loads will be highly variable in a way that depends on the relative thicknesses and
Fig. 28.18 'OSC' failures for combined bearing
bypass loading17.
Mechanically fastened joints 623 moduli of adjacent plate elements in the regions around each fastener; fastener deflections due to beam-bending of the fasteners as well as clearance effects will add other complications to the situation. In Fig. 28.19, configuration A illustrates the behavior of joints with both elements tapered (i.e. scarf joints); due to load transfer between the elements by fastener load, the net section load P, decreases from the loaded (thick) end of each plate to its unloaded (thin)end, while the corresponding thicknesses decrease keeps the gross section stresses and strains and therefore the stretching deflections uniform along each element, providing for nearly equal load transfer at each fastener. For configuration B of Fig. 28.19, the case of uniform element thickness, the interior fastener loads are smaller than those at the joint ends, which is typical of this situation. Since the configurations shown at the right in Fig. 28.19 are arranged in order of increasing load capability (see the strains
Fig. 28.20 Two fastener joint.
listed under each configuration), it is noted that the scarfed configuration gives the lowest strength of the four, and is about 9% weaker than the uniform thickness configuration, B. It is usually expected that scarfing will lead to a way of keeping the bearing load minimized at the joint ends where the highest P,s are encountered, but other factors having to do with the balance between the effects of local
BOLT LOAD DISTRIBUTIONS 4-ROW BOLTED NINT
-
CONFIGURATION D
E,
O11(#lNIIN.
CONFIGURATIONC
-
= 0A)OMINIIN.
E-
CONFIQURATION A CONFIQURATION B 0- 4 CONFIGURATIONC 1n CONFIGURATION D O----Q
-
CONFIGURATION B 0#)46INIIN. Ob0IN.
1
2 3 BOLT NUMBER
1
1 A 112. 112. 112
3 4 0251N.
!
i I
-
.'I
-
CONFIGURATION A E , O m 1 INAN.
Fig. 28.19 Effectof joint configuration on fastener load di~tribution'~.
-
2
4
I'
624 Mechanical fastening and adhesive bonding element thickness and local bearing load predominates here. Note that configuration D, the strongest, uses a combination of variable fastener diameter and element tapering to achieve a maximum thickness of the outer elements which is greater than that in configuration C, to obtain maximum joint strength. Thus, judicious use of element tapering and variation of fastener diameters as well as other joint parameters can improve joint performance. With untapered joints, the maximum benefit of additional rows of fasteners is not much more than 20% greater than that for single row joints. Joint tapering will provide some improvement over that figure, although the benefit is limited; it should be kept in mind that the net section load at the loaded end of a given element will be the same for single fastener and multi-fastener joints, so that the benefit of adding more than a few rows may not be great. The interaction of the effects encountered in multi-fastener joints is fairly complicated and requires the use of analyses which can take into account the stresses and strains in single fastener configurations with bypass loading present (Fig. 28.13), representing 'unit cells' of the joint configuration, together with finite element calculations which evaluate the interactions among the various unit cells to provide the overall fastener load distribution. Computer codes have been developed under DoD and NASA sponsorship to provide for this type of integrated joint design. For example, Nelson, Bunin and Hart-Smith13 discuss the application of the well-known 'A4EJ' codeI9in conjunction with the code 'BJSFM' (Bolted Joint Stress Field Model15)which were developed by McDonnell Douglas under NASA and Air Force sponsorship for this type of joint design analysis. Northrop similarly developed codes 'SASCJ' (Stress Analysis of Single Fastener Composite Joints) and 'SAMCJ' (Stress Analysis of Multifastener Composite Joints) under Air Force Contract6,20. For information on the theory and application of these codes, the reader is directed to the references.
28.3.5 FASTENER EFFECTS
Joint strengths for local areas (Fig. 28.13) around individual fasteners are affected by a number of parameters associated with the fasteners. Some of these include: whether the joint is in single or double shear (Fig. 28.21); the use of countersunk (flush head) compared with protruding head fasteners (Fig. 28.22); effects of fastener diameter; effects of fastener clamping; fastener clearance effects, and effect of fastener deflections on fastener load distribution. Bearing strengths are significantly affected by the use of single compared with
p7=g=7p2p
P
(A)
Double Shear Configuration
~
P I'2
(B) Single Shear configuration Fig. 28.21 Single shear and double shear configurations.
+fP
3 (A)
Protruding Head Fastener
P P
(B) Countersunk Fastener Fig. 28.22 Countersunk and protruding head fasteners.
Mechanicallyfastened joints 625 double shear configurations,bearing strengths for single shear joints tending to run considerably below those for double shear because of greater through-the-thickness variation of fastener-plate contact pressure. Bearing strength tests referenced in Section 28.3.6 include separate test configurations for the two situations. In addition, as indicated in Fig. 28.22(a),bending moments tend to occur in single shear joints which are not present with double shear arrangements. Fastener head pull through (Fig. 28.10) can be a problem in the presence of such bending effects, and special test methods for fastener pull-through strength are described in Section 28.3.6. Countersunk, or flush head, fasteners (Fig. 28.22(b)) are frequently encountered in exterior surfaces of aircraft components where avoidance of air flow disturbance is required. Countersunk fasteners for composites include (Fig. 28.23) 'tension head' fasteners having the larger head depths and therefore wider heads, and 'shear head' fasteners having smaller head depths, with head angles ranging from 100" to 130". Countersunk fasteners tend to bear against the surrounding element more unevenly through the thickness than protruding head fasteners do. Tension head fasteners are generally preferred over shear head fasteners because of greater strength against head pull-through; however, if the joint element is so thin that the countersunk depth is greater than 70% of the element thickness, the tendency toward uneven bearing pressure in tension head fasteners is too great and shear head fasteners are recommended in this case.
The fastener diameter should be on the order of the thickness of the thicker of the plate elements making up the joint, or greater, ( D / t 2 1) to avoid excessive fastener bending. As in Fig. 28.10 (lower right-hand sketch) excessive bending can lead to failure of the fasteners, which is intolerable. In addition, fastener bending causes uneven distribution of bearing pressure through the plate element thicknesses, so that the full bearing strength is not available in such cases. The effect of large fastener deflections on the clamping pressure provided by the fastener is another adverse effect of fastener bending deflections. Figure 28.2413 illustrates the fact that bending deformations reduce the clamping pressure provided by fastener head, causing a reduction of bearing strength which is in addition to that caused by uneven bearing pressure through the thickness. The beneficial effect of clamping pressure on bearing strength, discussed earlier, has been well established. Required clamping levels are usually described in terms of bolt torques, 'finger tightened' being the lightest level, and installation requirements specify torque levels which supposedly represent particular bolt tensions (and therefore clamping
TENSION JOINT HlQH BEARING LOAD SIDE
LOELAMINATIONS DUE TO BEARING LOAD AND REDUCED HlQH BEARINQ LOAD SIDE
COMPRESSION JOINT
w-
Fig. 28.24 Effects of fastener bending on joint fail-
Fig. 28.23 Tension head and shear head fasteners.
ureI3.
626 Mechanical fastening and adhesive bonding pressures) that can be calculated in terms of the pitch of the fastener threads from machine design formulas. Bolt tensions for a given torque level are notoriously variable because of friction effects in the bolt threads, but specified torque levels which have been determined empirically probably represent minimum clamping levels necessary to insure maximum bearing strengths that can be achieved when variations in service conditions are taken into account. Steps should be taken to avoid loss of clamping pressure due to through-the-thickness viscoelastic deformation of the laminatez1”at elevated temperature and humidity. Fastener clearances are typically on the order of 0.075 mm (0.003 in) or less for typical 0.635 mm (0.25 in) aircraft fasteners. Analytical studies have shown that bearing stresses increase significantly for such relatively small clearances since the angle of contract decreases rapidly as the clearance increases. Clearances also have a significant effect on the fastener load distribution since the last in a series of fasteners cannot take up load until all the clearances have been taken up. In addition to the effect of clearances on fastener load distribution, effects of fastener bending deflections must be taken into account in load distribution calculations such as those provided by the A4EJ, SASCJ and SAMCJ codes described above. In the case of the two-fastener joint shown in Fig. 28.20, bending and rotational deflections of the fasteners will modify the load distribution described in the discussion of that figure for in Fig. zero fastener deflection. For E,t,>>E,t,, 28.20 for example, fastener deflections will allow some load to be transferred to the second fastener, as opposed to the case of no fastener deflections discussed earlier which led to Pf2= 0. Fastener deflection effects can be inferred from bolt bearing tests which provide for deflection measurements. Alternatively, analytical approaches based on beam models for the fastener which include both bending and shear deformations have been used13.
Complications, such as the way in which the fastener head and nut/washer combination bears on the surfaces of the plate element will influence the outcome of such calculations and must be taken into account. In addition, the through-thickness distribution of bearing pressure between the fastener and the surrounding the plate element should be included in the calculations. The method of Harris, qalvo and Hoosonz3which treats the bore of the fastener hole as an elastic foundation for the beam used to model the fastener has been applied6for such calculations. 28.3.6 TEST METHODS
Joint strength tests are needed to establish certain key parameters of the joint as inputs to design analyses. Such data as failure stresses for pure bearing load as obtained from single fastener coupons, and open hole coupon strengths for both tension and compression loading, are needed to establish joint performance for pure bearing and bypass loads. Intermediate combinations of bearing and bypass load must also be considered to provide empirical curves for dealing with the general situation. Because of differences in the way the fastener contacts the surrounding plate materials, bearing tests have to be conducted to treat both single and double lap (single and double shear) configurations. In single lap joints in particular, tests are needed to establish the effect of fastener rotation and bending deflection. Fastener deflections must be determined in tests of the type just described for providing fastener response data in connection with predictions of load distribution in multifastener joints. In addition, fastener head pull-through strength tests have to be performed to allow for joint configurations in which overall bending takes place, in which case out-of-plane forces between the fastener and joint plates tend to be sigruficant. The details of test methods for mechanically fastened joints are described by Shyprykevichz4 and in Mn-HDBK-1725.
Adhesive joints 627 28.4 ADHESIVE JOINTS
28.4.1 INTRODUCTION
Adhesive joints are capable of high structural efficiency and constitute a resource for structural weight saving because of the potential for elimination of stress concentrations which cannot be achieved with mechanically fastened joints. Unfortunately, because of a lack of reliable inspection methods and a requirement for close dimensional tolerances in fabrication, aircraft designers have generally avoided bonded construction in primary structure. Some notable exceptions include: bonded step lap joints used in attachments for the F-14 and F-15 horizontal stabilizers as well as the F-18 wing root fitting, and a majority of the airframe components of the Lear Fan and the Beech Starship. While a number of issues related to adhesive joint design were considered in the earlier literaturezG3, much of the methodology currently used in the design and analysis of adhesive joints in composite structures is based on the approaches evolved by L.J. HartSmith in a series of NASA/Langley-sponsored contracts of the early 1 9 7 0 as~ well ~ ~as~from the Air Force’s Primary Adhesively Bonded Structures Technology (PABST) programw3 of the mid-1970s. The most recent such work developed three computer codes for bonded and bolted joints, designated ‘A4EG’, ’A4EI’ and ’A4EKW under Air Force Contract . The results of these efforts have also appeared in a number of open literature publi~ations~’-~. In addition, such approaches found application in some of the efforts taking place under the NASA Advanced Composite Energy Efficient Aircraft (ACEE) program of the early to mid198Os5O,51. Some of the key principles on which these efforts were based include: (1)the use of simple one-dimensionalstress analyses of generic composite joints wherever possible; ( 2 ) the need to select the joint design so as to ensure failure in the adherend rather than the adhesive, so that
the adhesive is never the weak link;(3) recognition that the ductility of aerospace adhesives is beneficial in reducing stress peaks in the adhesive; (4) careful use of such factors as adherend tapering to reduce or eliminate peel stresses from the joint; (5) recognition of slow cyclic loading, corresponding to such phenomena as cabin pressurization in aircraft, as a major factor controlling durability of adhesive joints, and the need to avoid the worst effects of this type of loading by providing sufficient overlap length to ensure that some of the adhesive is so lightly loaded that creep cannot occur there, under the most severe extremes of humidity and temperature for which the component is to be used. Much of the discussion to follow will retain the analysis philosophy of Hart-Smith, since it is considered to represent a major contribution to practical bonded joint design in both composite and metallic structures. On the other hand, some modifications are introduced here. For example, the revisions of the Goland-Reissner single lap joint analysis36 have been re-revised according to the approach presented in Refs. 53,54. Certain issues which are specific to composite adherends but were not dealt with in the Hart-Smith efforts will be addressed. The most important of these is the effect of transverse shear deformations in organic composite adherends. 28.4.2 SUMMARY OF JOINT DESIGN CONSIDERATIONS
28.4.2.1 Effects of adherend thickness: adherend failures versus bond failures Figure 28.25 shows a series of typical bonded joint configurations. Adhesive joints in general are characterized by high stress concentrations in the adhesive layer. These originate, in the case of shear stresses, because of unequal axial straining of the adherends, and in the case of peel stresses, because of eccentricity in the load path. Considerable ductility is associated
628 Mechanical fastening and adhesive bonding
-
1
- (0) - 4
TAPERED SINGLE-UP JOINT DOUBLE-UPJOINT (F)
1 DUJBLE-STRAP JOINT
(0)
illustrate this point, shows a progression of joint types which represent increasing strength capability from the lowest to the highest in the figure. In each type of joint, the adherend thickness may be increased as an approach to achieving higher load capacity. When the adherends are relatively thin, results of stress analyses show that for all of the joint types in Fig. 28.26, the stresses in the bond will be small enough to guarantee that the adherends will reach their load capacity before failure can occur in the bond. As the adherend thicknesses increase, the bond stresses become relatively larger until a point is reached at which bond failure occurs at a lower load than that for which the adherends fail. This leads to the general principle that for a given joint type, the adherend thicknesses should be restricted to an appropriate range relative to the bond layer thickness. Because of processing considerations and defect sensitivity of the bond material, bond layer thicknesses are generally limited to a range of 0.125-0.39 mm (0.005-0.015 in). As a result, each of the joint
n
4
f
&
TAPERED STRAP JOINT
Fig. 28.25 Adhesive joint types”,
55.
with shear response of typical adhesives, which is beneficial in minimizing the effect of shear stress joint strength. The response of typical adhesives to peel stresses tends to be much more brittle than that to shear stresses, and reduction of peel stresses is desirable for achieving good joint performance. From the standpoint of joint reliability, it is vital to avoid the condition where the adhesive layer is the weak link in the joint, i.e. that the joint be designed to ensure that the adherends fail before the bond layer whenever possible. T ~ isE because failure in the adherends may be controlled, while failure in the adhesive is resin dominated, and thus subject to effects of voids and other defects, thickness variations, environmental effects, processing variations, deficiencies in surface preparation and other factors that are not always adequately controlled. This is a significant challenge, since adhesives are inherently much weaker than the composite or metallic elements being joined. However, the objective can be accomplished by recognizing the limitations of the joint geometry being considered and placing appropriate restrictions on the thicknesses the adherends for any given geometry. Figure 28.26, which 55 to has frequently been used by Hart-Smith39,
ADHEREND THICKNESS
Fig. 28.26 Joint geometry effects39.
Adhesive joints 629 types in Figs. 28.25 and 28.26 corresponds to a specific range of adherend thicknesses and therefore of load capacity, and as the need for greater load capacity arises, it is preferable to change the joint configuration to one of higher efficiency rather than to increasing the adherend thickness indefinitely. 28.4.2.2 Joint geometry effects Single and double lap joints with uniformly thick adherends (Fig. 28.25(b), (e) and ( f ) ) are the least efficient joint type and are suitable primarily for thin structures with low running loads (load per unit width, i.e. stress times element thickness). Of these, single lap joints are the least capable because the eccentricity of this type of geometry generates significant bending of the adherends that magnifies the peel stresses. Peel stresses are also present in the case of symmetric double lap and double strap joints, and become a limiting factor on joint performance when the adherends are relatively thick. Tapering of the adherends (Figs. 28.25(d) and (g)) can be used to eliminate peel stresses in areas of the joint where the peel stresses are tensile, which is the case of primary concern. For joints between adherends of identical stiffness, scarf joints (Fig. 28.25(i))are theoretically the most efficient, having the potential for complete elimination of stress concentrations. (In practice, some minimum thickness corresponding to one or two ply thicknesses must be incorporated at the thin end of the scarfed adherend leading to the occurrence of stress concentrations in these areas.) In theory, any desirable load capability can be achieved in the scarf joint by making the joint long enough and thick enough. However, practical scarf joints may be less durable because of a tendency toward creep failure associated with a uniform distribution of shear stress along the length of the joint unless care is taken to avoid letting the adhesive be stressed into the nonlinear range. As a result, scarf joints tend to be used only for repairs of very thin structures.
Scarfjoints with unbalanced stiffnessesbetween the adherends do not achieve the uniform shear stress condition of those with balanced adherends, and are somewhat less structurally efficient because of rapid buildup of load near the thin end of the thicker adherend. Step lap joints (Fig. 28.25(h)) represent a practical solution to the challenge of bonding thick members. This type of joint provides manufacturing convenience by accommodating the layered structure of composite laminates. In addition, high loads can be transferred if sufficiently many short steps of sufficiently small ’rise’ (i.e. thickness increment) in each step are used, while maintaining sufficient overall length of the joint. 28.4.2.3 Effects of adherend stiffness unbalance All types of joint geometry are adversely affected by unequal adherend stiffnesses, where stiffness is defined as axial or in-plane shear modulus times adherend thickness. Where possible, the stiffnesses should be kept approximately equal. For example, for step lap and scarf joints between quasi-isotropic carbon epoxy (Young’s modulus = 55 GPa = 8 x lo6 lb/in2) and titanium (Young’s modulus = 110 GPa = 16 x lo6 lb/in2) ideally, the ratio of the maximum thickness (the thickness just beyond the end of the joint) of the composite adherend to that of the titanium should be 110/55 = 2.0. 28.4.2.4 Effects of ductile adhesive response Adhesive ductility is an important factor in minimizing the adverse effects of shear and peel stress peaks in the bond layer. If peel stresses can be eliminated from consideration by such approaches as adherend tapering, strain energy to failure of the adhesive in shear has been shown by Ha~?-Smith~~ to be the key parameter controlling joint strength; thus the square root of the adhesive strain energy
630 Mechanical fastening and adhesive bonding density to failure determines the maximum sta- transverse tension, as a result of which the tic load that can be applied to the joint. The limiting element in the joint may be the interwork of Hart-Smith has also shown that for pre- laminar shear and transverse tensile strengths dicting mechanical response of the joint, the of the adherend rather than the bond strength. detailed stress-strain curve of the adhesive can Ductile behavior of the adherend matrix can be replaced by an equivalent curve consisting be expected to have an effect similar to that of of a linear rise followed by a constant stress ductility in the adhesive in terms of response plateau (i.e. elastic-perfectly plastic response) if of the adherends to transverse shear stresses, the latter is adjusted to provide the same strain although the presence of the fibers probably energy density to failure as the actual limits this effect to some extent, particularly in stress-strain curve gives. Test methods for regard to peel stresses. The effect of the stacking sequence of the adhesives should be aimed at providing data on this parameter. Once the equivalent elastic- laminates making up the adherends in composperfectly plastic stress-strain curve has been ite joints is sigruficant. For example, 90” layers identified for the selected adhesive for the most placed adjacent to the bond layer theoretically severe environmental conditions (temperature act largely as additional thicknesses of bond and humidity) of interest, the joint design can material, leading to lower peak stresses, while proceed through the use of relatively simple 0” layers next to the bond layer give stiffer one-dimensional stress analysis, thus avoiding adherend response with higher stress peaks. In the need for elaborate finite element calcula- practice it has been observed that 90” layers tions. Even the most complicated of joints, the next to the bond layer tend to seriously weaken step lap joints designed for root-end wing and the joint because of transverse cracking which tail connections for the F-18 and other aircraft, develops in those layers, and advantage cannot have been successfully d e ~ i g n e d ~ and ~ , ~ ”be ~ taken of the reduced stresses. Large disparity of thermal expansion charexperimentally demonstrated using such approaches. Design procedures for such analy- acteristics between metal and composite ses which were developed on Government adherends can pose severe problems. contract have been incorporated into public Adhesives with high curing temperatures may domain in the form of the ’A4EG’, ‘A4EI’ and be unsuitable for some uses below room tem‘A4EK computer codesmentioned previ- perature because of large thermal stresses ously in Section 28.4.1.Note that the A4EK code which develop as the joint cools below the fabpermits analysis of bonded joints in which local rication temperature. Composite adherends are relatively pervidisbonds are repaired by mechanical fasteners. ous to moisture, which is not true of metal adherends. As a result, moisture is more likely 28.4.2.5 Behavior of composite adherends to be found over wide regions of the adhesive Organic matrix composite adherends are con- layer, as opposed to confinement near the siderably more affected by interlaminar shear exposed edges of the joint in the case of metal and tensile stresses than metals, so that there is adherends, and response of the adhesive to a significant need to account for such effects in moisture may be an even more significant stress analyses of joints. Transverse shear and issue for composite joints than for joints thickness-normal deformations of the between metallic adherends. adherends have an effect analogous to thickening of the bond layer, corresponding to a 28.4.2.6 Effects of bond defects lowering of both shear and peel stress peaks. On the other hand, the adherend matrix is Defects in adhesive joints which are of concern often weaker than the adhesive in shear and include surface preparation deficiencies, voids
Adhesive joints 631 and porosity, and thickness variations in the adherends, porosity may grow catastrophibond layer. cally and lead to non-damage tolerant joint Of the various defects which are of interest, performance. surface preparation deficiencies are probably Bond thickness variations'jl usually take the the greatest concern. These are particularly form of thinning due to excess resin bleed at troublesome because there are no current non- the joint edges, leading to overstressing of the destructive evaluation techniques which can adhesive in the vicinity of the edges. Inside detect low interfacial strength between the tapering of the adherends at the joint edges bond and the adherends. Most joint design will compensate for this condition; other comprinciples are academic if the adhesion pensating techniques are also discussed'jl. between the adherends and bond layer is poor. Bond thicknesses, per se, should be limited to The principles for achieving good adhesion of ranges of 0.12-0.24 mm (0.005-0.01 in) to prethe bond to the adherends (see Chapter 29) are vent significant porosity from developing well established for adherend and adhesive although greater thicknesses may be acceptcombinations of interest. Hart-Smith, Brown able if full periphery damming or high and Won$ give an account of the most crucial minimum viscosity paste adhesives are used. features of the surface preparation process. Common practice involves the use of film Results shown in that reference suggest that adhesives containing scrim cloth, some forms surface preparation which is limited to of which help to maintain bond thicknesses. It removal of the peel ply from the adherends is also common practice to use mat carriers of may be suspect, since some peel plies leave a chopped fibers to prevent a direct path for residue on the bonding surfaces that makes access by moisture to the interior of the bond. adhesion poor. (However, some manufacturers have reported satisfactory results from 28.4.2.7 Durability of adhesive joints surface preparation consisting only of peel ply removal.) Low pressure grit b l a ~ t i n g ~ is ~ ,Hart-Smith45 ~~ discusses differences in durabilpreferable over hand sanding as a means of ity assessment of adhesive joints between eliminating such residues and mechanically concepts related to creep failure under cyclic conditioning the bonding surfaces. loading and those related to crack initiation For joints which are designed to ensure that and propagation which require fracture the adherends rather than the bond layer are mechanics approaches for their interpretation. the critical elements, tolerance to the presence In summary, Hart-Smith suggests that if peel of porosity and other types of defect is consid- stresses are eliminated by adherend tapering erable45.Porosity'jOis usually associated with or other means, and if the principle discussed overthickened areas of the bond, which tend in Section 28.4.2.1 of limiting the adherend to occur away from the edges of the joint thickness to ensure failure of the adherends where most of the load transfer takes place, rather than the adhesive is followed, crackand thus is a relatively benign effect, espe- type failures will not be observed under cially if peel stresses are minimized by time-varying loading, failures being related adherend tapering. In such cases6", porosity primarily to creep fatigue at hot wet condican be represented by a modification of the tions, in joints with short overlaps which are assumed stress-strain properties of the adhe- subject to relatively uniform distributions of sive as determined from thick-adherend tests, shear stress along the joint length. Additional allowing a straightforward analysis of the discussion of viscoelastic response of bonded effect of such porosity on joint strength, as in joints is There is an extensive body of literature6571 the A4EI computer code. If peel stresses are significant, as in the case of over-thick on fracture mechanics approaches to joint
632 Mechanical fastening and adhesive bonding durability, based on measurement of energy release rates for various adhesives together with analytical efforts aimed at applying them to joint configurations of interest. In particular, Johnson and Malln report fatigue crack initiation in bonded specimen configurations with adherend tapering aimed at reduction of peel stresses in varying degrees, in some cases practically eliminating them; data in Ref. 92 indicate that crack initiation will occur even with the adhesive in pure shear, for cycling to lo6cycles above loading levels which are probably considerably below static failure loads. The results given” suggest that for combinations of peel and shear stressing, total (mode 1 + mode 2) cyclic energy release rate can be used to determine whether or not cracking will occur. However, Hart-Smith reportedfi that in ’thick adherend‘ test specimens that provide a relatively uniform shear stress distribution in the adhesive (see MIL-HDBK-17, Vol. 1, Chapter 7, Section 7.3) which were subjected to fatigue tests in the PABST programM,cycling to more than lo7 cycles applied at high cycling rates (30 Hz) were achieved without failure of the adhesive, although in certain cases, namely those involving 6.27 mm (0.25 in) adherend thicknesses, fatigue failures of the metal adherends did result. More study is needed to resolve some of the apparently contradictory results which have come out of various studies. 28.4.3 STRESSES IN ADHESIVE JOINTS
28.4.3.1 General
Stress analyses of adhesive joints have ranged from very simplistic ‘P over A’ formulations in which only average shear stresses in the bond layer are considered, to extremely elegant elasticity approaches that consider fine details, e.g. the calculation of stress singularities for application of fracture mechanics concepts. A compromise between these two extremes is desirable, since the design of structural joints does not usually depend on the fine details of the stress distributions. Since practical consid-
erations force bonded joints to incorporate adherends which are thin relative to their dimensions in the load direction, stress variations through the thickness of the adherend and the adhesive layer tend to be moderate. Such variations do tend to be more sigruficant for organic composite adherends because of their relative softness with respect to transverse shear and thickness normal stresses. However, a considerable body of design procedure has been developed based on ignoring thicknesswise adherend stress variations. Such approaches involve using one-dimensional models in which only variations in the axial direction are accounted for. Accordingly, the bulk of the material to be covered here is based on simplified one-dimensional approaches characterized by the work of Hart-Smith. The Hart-Smith approach makes extensive use of closed form and classical series solutions since these are ideally suited for making parametric studies of joint designs. The most prominent of these have involved modification of Volkersen26 and Goland-ReissnerZ7solutions to deal with ductile response of adhesives in joints with uniform adherend thicknesses along their lengths, together with classical series expressions to deal with variable adherend thicknesses encountered with tapered adherends, and scarf joints. Simple lap joint solutions described below calculate shear stresses in the adhesive for various stiffnesses and applied loadings. For the more practical step lap joints, the described expressions can be adapted to treat the joint as a series of separate joints, each having uniform adherend thickness. 28.4.3.2 Adhesive shear stresses
Figure 28.27 shows a joint with ideally rigid adherends in which neighboring points on the upper and lower adherends slide horizontally with respect to each other when the joint is loaded to cause a displacement difference 6 = uu - uLrelated to the bond layer shear strain by yb = 6 / f b .The corresponding shear stress, zb, is given by zb = Gbyb. The rigid adherend
Adhesive joints 633 one for which E,tL >> E&), stretching elongations in the upper adherend lead to a shear strain increase at the right end of the bond layer. The case in which both adherends are equally deformable, shown in Fig. 28.29(b), indicates a bond shear strain increase at both ends due to the increased axial strain in dTU/dx= zb (28.6) whichever adherend is stressed at the end leads to a linear distribution of Tu and TL under consideration. For both cases, the varia(upper and lower adherend resultants) as well tion of shear strain along the bond results in an as the adherend axial stresses uxuand ax,indi- accompanying increase in shear stress which, cated in Fig. 28.28. These distributions are when inserted into the equilibrium eqn (28.6) leads to a nonlinear variation of stresses. The described by the following expressions: Volkersen shear lag analysisz6provides the simplest calculation of adhesive shear stresses for the case of deformable adherends. This involves the solution of the following differential equation:
assumption implies that 6, y, and t, are uniform along the joint. Furthermore, the equilibrium relationship indicated in Fig. 28.27(c),which requires that the shear stress be related to the resultant distribution in the upper adherend by
where ax = T/t. In actual joints, adherend deformations will cause shear strain variations in the bond layer which are illustrated in Fig. 28.29. For the case of a deformable upper adherend in combination with a rigid lower adherend shown in Fig. 28.29(a) (in practice,
B,
E&;
B, = E,t,
(28.8)
which applies to the geometry of Fig. 28.30
I [a RIGID
=
KkU
f-
Fig. 28.27 Elementary joint analysis (rigid adherend model).
I
634 Mechanical fastening and adhesive bonding (A)
(a) AXIAL STRESS DISTRIBUTXOH
AXIAL RESULTANT DISTRIBUTION
Fig. 28.28 Axial stresses in joint with rigid adherends.
[A) RIGID UIIW
AaEml
--f
TFig. 28.29 Adherend deformations in idealized joints.
below. The solution for this equation which provides zero traction conditions at the left end of the upper adherend and the right end of the lower adherend, together with the applied load T at the loaded ends gives the resultants as:
-
TL = T - T U
where
-
t=-
+-
1 +PB SinhPZ/t
(28.9)
tu + tL ; PB = 2
BL/B"
Using eqn (28.6) to obtain an expression for the shear stress distribution leads to:
Adhesive joints 635 I 0I
X
+- PB
1 + pB tanhpZ/T
)
(28.11)
where Gx= Fig. 28.30 Geometry for Volkersen solution.
Also of interest in the discussion which follows is the minimum shear stress in the joint. This occurs approximately at x = 1/2, leading to:
B, 2 B,;
I
0
44
&4
a6
Q1
T/t
to be discussed subskquently. Figure 28.31
t
1.2
1.4
1s
1 1
2
IS
1
I
.-. .
4
U a.4
.
.
.
:
:
.
.
OS
a.0
1
IS!
1.4
IJ
1.8
#-
b
2
. . . . . . . . . . o)
a4 w os
1
IZ 11) i g t i
2
I---
I
636 Mechanicalfastening and adhesive bonding
Fig. 28.32 Comparison of average and maximum shear stress vs. l / t .
shows the distribution of axial adherend stresses and bond layer shear stress for two cases corresponding to E, = E , and E, = 10Eu with tu = t,, p = 0.387 and l / t = 20 for both cases (giving p l / t = 7.74) and a nominal adherend stress 0, = 10. As in the approximate analysis given earlier, the shear stresses given by eqn (28.10) are maximum at both ends for equally deformable adherends (B, = B,); for dissimilar adherends with the lower adherend more rigid (B, > E$,), the maximum shear stress obtained from eqn (28.10) occurs at the right end of the joint where x = I , again as it did for the approximate analysis. Figure 28.32 compares the behavior of the maximum shear stress with the average shear stress as a function of the dimensionless joint length, l / t , for equal adherend stiffnesses. The point illustrated here is the fact that although the average shear stress continuously decreases as the joint length increases, for the maximum shear stress which controls the load that can be applied without failure of the adhesive, there is a diminishing effect of increased joint length when q = p l / t is much greater than about 2.
An additional point of interest is a typical feature of bonded joints illustrated in Fig. 28.31(d) which gives the shear stress distribution for equal adherend stiffness, namely, the fact that high adhesive shear stresses are concentrated near the ends of the joint. Much of the joint length is subjected to relatively low levels of shear stress, which implies in a sense that that region of the joint is structurally inefficient since it does not provide much load transfer. However, the region of low stress helps to improve damage tolerance of the joint since defects such as voids and weak bond strength may be tolerated in regions where the shear stresses are low, and in joints with long overlaps this may include most of the joint. In addition, Hart-Smith has suggested51 that when ductility and creep are taken into account, it is a good idea to have a minimum shear stress level no more than 10% of the yield strength of the adhesive, which requires some minimum value of overlap length. Equation (28.12) can be used to satisfy this requirement for the case of equal stiffness adherends. The two special cases of interest again are for equal adherend stiffness and a
Adhesive joints 637 rigid lower adherend, since these bound the range of behavior of the shear stresses. As a practical consideration, we will be interested primarily in long joints for which pZ/t >> 1. For these cases eqn (28.11) reduces to:
relatively obvious due to the offset of the two adherends which leads to bending deflection as in Fig. 28.33@).In the case of double lap joints, as exemplified by the configuration shown in Fig. 28.34, the load path eccentricity is not as obvious, and there may be a tendency p1/t >> 1; to assume that peel stresses are not present for this type of joint because, as a result of the lateral symmetry, there is no overall bending deflection. However, a little reflection brings to mind the fact that while the load in the sym1 B, = B,; zJrnaX=-pax (28.13) metric lap joint flows axially through the 2 central adherend prior to reaching the overlap Thus, for long overlaps, the maximum shear region, there it splits in two directions, flowing stress for the rigid adherend case tends to be laterally through the action of bond shear twice as great as that for the case of equally stresses to the two outer adherends. Thus deformable adherends, again illustrating the eccentricity of the load path is also present in adverse effect of adherend unbalance on shear this type of joint. As seen in Fig. 28.34(c), the shear force, designated as F,, which represtress peaks. sents the accumulated effect of zb for one end of the joint, produces a component of the total 28.4.3.3 Peel stresses moment about the neutral axis of the upper Peel stresses, i.e. through-the-thickness exten- adherend equal to FsHz/2. (Note that F , is sional stresses in the bond, are present because equivalent to T/2, since the shear stresses react the load path in most adhesive joint geome- this amount of load at each end.) The peel tries is eccentric. It is useful to compare the stresses, which are equivalent to the forces in effect of peel stresses in single and double lap the restraining springs shown in Fig. 28.34(b) joints with uniform adherend thickness, since peel stresses are most severe for joints with uniform adherend thickness. The load path eccentricity in the single lap joint (Fig. 28.33) is
, \
i Fig. 28.33 Peel stress development in single lap ioints.
Fig. 28.34 Peel stress development in double lap ioints.
J
638 Mechanical fastening and adhesive bonding It is important to understand that peel and (c) have to be present to react the moment stresses are unavoidable in most bonded joint produced by the offset of FsHabout the neutral configurations However, they can often be axis of the outer adhered. Peel stresses are reduced to acceptable levels by selecting the highly objectionable. Later discussion will adherend geometry appropriately. indicate that effects of ductility significantly reduce the tendency for failure associated with shear stresses in the adhesive. On the other 28.4.3.4 Effects of joint geometry hand, the adherends tend to prevent lateral contraction in the in-plane direction when the In this section the behavior of joints is considbond is strained in the thickness direction, ered with linear response of the adhesive in which minimizes the availability of ductility shear assumed. Effects of ductility will be coneffects that could provide the same reduction sidered later. of adverse effects for the peel stresses. This is illustrated by the butt tensile test shown in Fig. 28.4.3.4.1 Single and double lap joints with 28.35 in which the two adherend surfaces adjauniform adherend thickness cent to the bond are pulled away from each other uniformly. Here the shear stresses asso- Double lap joints will be considered first since ciated with yielding are restricted to a small they are somewhat simpler to discuss than sinregion whose width is about equal to the gle lap joints because of deflection effects in the thickness of the bond layer, near the outer latter. Shear and peel stresses in double lap edges of the system; in most of the bond, rela- joints with uniform adherend thickness were tively little yielding can take place. For organic treated by Hart-Smith%.For the shear stresses, matrix composite adherends, the adherends the type of analysis discussed in Section 28.4.3.2 may fail at a lower peel stress level than that at can be applied with suitable changes in notawhich the bond fails, which makes the peel tion, i.e. the expressions for the shear stresses given in eqns (28.11)and (28.12)can be applied stresses even more undesirable. with subscripts 'i' and '0' ('inner' and 'outer') substituted here for 'L' and ' U ('lower' and 'upper') used in eqns (28.6-28.11); in addition, the outer adherend thickness in the earlier equations is now equivalent to half the thickness of inner adherend because of vertical symmetry of the double lap joint. However, we will also introduce the effects of thermal mismatch effects in the following expressions for later reference. The notation used here is: Bond
P t
P
Bo = toEo; Bi = tiEi; E@e Region (Distornal Strains)
pB = Bi/Bo;
B B. Tth = (ao- a,)AT; Bo + Bi L A
Fig. 28.35 Shear stresses near outer edges of butt tensile test.
ax = T / t ; &* =T,/t
-
(28.14)
Adhesive joints 639 where a,, a, are thermal expansion coefficients and AT is the temperature change. Note that is related to the resultants (axial adherend stress times thickness) at the ends of the joint as shown in Fig. 28.36. The shear stresses are then given by: Zb
[
=/35
coshp(x - I)/: +pB sinhBl/t 1
~
assuming that Bi 1 Bo, the maximum value of the shear stresses occurs at the right end of the joint as noted earlier (Fig. 28.31). With thermal effects present, the situation is complicated by the sign of &* which is positive if (a, - a,)and AT have the same sign and negative otherwise. The peel stresses in the double lap joint are described by a beam-on-elastic foundation type differential equation of the form:
**--
-
coshp(Z - x ) / t
]
(A) DOUBLE STRAP JOINT
2t
1*’ 1
d40 ?d b + 4 - 0 dP t4
=-f-
2
yd
=
O 114
(28.18b)
The solution to eqn (28.18) depends on whether a strap joint or a lap joint is considered. The exact form of the solution contains products of hyperbolic and trigonometric functions but for the practical situation of joints longer than one-or-two adherend thicknesses and B << yd, are given by:
WUBLE LAP JOINT
Double strap joint,
For the usual situation in which the overlap is long enough so that p l / t is greater than about 3, the peak shear stresses at the ends of the joint are given by: = 0;
Zb,
=
B(&
(28.19) For the case of identical adherends, the maximum peel stresses, which occur at x = 0, are given by:
5,- &&
Ob)max
and for the special case of equal adherend stiffnesses (Bi = Bo) we have: Bi = BJp, = 1);
‘b)-x
=
1
(28.18a)
Double lap joint, (E)
Fig. 28.36 Symmetric double strap/double lap joints.
x
dzb dx
(3x) Ebto
(28.15)
sinh pz /t
2Tt
h
In the absence of thermal effects ( Tth = 0) and
T p ax “*
‘b)-x
=
= ‘b)max yd
(28.20)
P ax/2 - p
(identical adherends)
Here z ~is taken ) ~to be~ the peak stress at the left end of the joint, corresponding to the expression for x = 0 in eqn (28.16), since the out-of-plane normal stresses are compressive (28*17) at the other end of the joint for a tensile load.
640 Mechanicalfastening and adhesive bonding For compressive loading, the situation would Quarter Plane Symmetry reverse for the double lap joint (Fig. 28.36(b)), with the positive out-of-plane stresses occurring at the right end of the joint ( x = l), in the /t b case of the double snap joint (Fig. 28.27(a)),the peak out-of-plane stresses would be compressive at the left end of the joint and would not occur at x = 1, since the inner adherends butt against each other there and act as a continuous element. I 4 Effects of thermally induced stresses will be discussed in a later section. Figure 28.37 com2.00 3 pares the peel and shear stress distributions for 8, = 0, in a typical joint having balanced adherend stiffnesses (the sum of the outer adherend stiffnesses equal to the inner adherend stiffness) whose parameters are listed in Fig. 28.37(b). The diagram at the top indicates the origin of x at the left end of the overlap. The distribution of peel stresses is ~x-,-,,A0.00. : : : : : : somewhat more concentrated near the ends 0.00 0.200.400.60 0.801 .OO 1.20 1.40 1.601.80 2.00 than that of the shear stresses and the peel (a) Xstresses at the right end of the joint are negative. In addition, the compressive peak at the right end is half as great for the strap joint as for the lap joint, which is the result of the restraint of bending rotations in the strap joint m for a gap which is essentially zero. If the loadg o ing were compressive rather than tensile, the %-0.5 Xinner adherends would bear directly on each other and no shear or peel stress peak would Double strap joint occur at the gap, whereas in the lap joint the -1.5 right end of the overlap would experience the (b) same peak stresses for compressive loading as the left end does for tensile loading. Fig. 28.37 Bond stresses in double lap/strap joints; The situation for the single lap joint (Fig. (a) bond shear stress distribution; (b) bond peel 28.38) is complicated by the effects of lateral stress distribution. deflection which are indicated in Fig. 28.39. Literature for the following discussion on the The effects of lateral deflections on the bond single lap joint is given53,51,7*78. stresses were first evaluated by The deflection effect is dependent on the Goland-ReissnerZ7 for the case of equal joint load, given in terms of the quantity adherend thicknesses, so that tu and t, can be LIl/2(8)1/2tu, where denoted by t in the following. The lateral deflections can then be stated in terms of a (28.21) dimensionless ratio, k, with respect to the LI = tud( E 12%); D, = EEutA 1 adherend thickness, and are of the following form:
t
t
t
p-'
i
v)
v)
c){(
Adhesive joints 641
- I -E
wL" DI PLACEMENTS
Fig. 28.38 Single lap joint geometry.
Fig. 28.39 Effects of bending deflections in single lap joints.
effects. The most accurate expression (given in Ref. 74) is fairly elaborate and will not be repeated here; an expression of intermediate accuracygiven in Ref. 74 retains the essential form of the GR result but gives considerable improvement over the GR expression for thin adherends: tanhLUo k = (28.23a) tanhLUo+ d8C,tanh(LU/2Cp) where
t" = t, = t;
l o ~ x < l + l o ;= w + -tu ---
2
2
U/.IS(X - L M ] w,sinh[ sinh[Ulfl8t] t 2
+ tb x ; W, = -(1 - k)
L
(28.22)
The Goland and Reissner (GR) expression for the parameter k has been re-examined by Hart-Smith%and more recently by Oplinger", based on the discussion in Ref. 74, the Goland-Reissner expression appears to provide adequate accuracy unless the adherends are excessively thin,not more that one or two times the bond layer thickness, in which case the expressions given in Refs. 73, 74 and provide corrections for bond layer thickness
The original GR expression for k is recovered if Cpis set equal to 1corresponding to tu >> t, (i.e. relatively thick adherends) and tanh LUo is likewise set to 1 corresponding to very long outer adherend lengths. A plot of k against the adherend loading stress is is given in Fig. 28.40 for two different values of adherend thickness corresponding to bond thickness-to-adherend thickness ratios (p, in eqn (28.23b)) of 0.5 and 0.1. This plot suggests that k is fairly constant at a value of about 0.25 for a wide range of applied stress values once the initial drop has occurred. The effect of bond-to-adherend thickness ratio is not particularly great and can perhaps be ignored for the most part.
642 Mechanical fastening and adhesive bonding 1
a9
ae
I
a7 0.6
* a5 0.4
a3
0.2
ai 0
Fig. 28.40 k parameter vs. adherend loading stress.
The lateral deflections of the joint have a significant influence on the stresses in the bond layer, which show this through the presence of the k parameter in expressions for them. The shear stress is given by: Zb =
+
-
'Osh
iu('
ab =a,- b
-
t - sin y
"-") ,
t
lo - x
where y, = (6 Ebt/Extb)1/4
(28.25)
L
lo+l-x ([YS(COS Y, t
,
- ')/.lst1
sinh ( d / 2 & )
s
x - lo
x - lo
where B and U are given in eqns (28.14) and (28.21). Equation 28.24, which represents a slight modification of the GR expression, reduces to the latter for small values of LIl/t. In addition, the peel stresses, for joints in which the overlap length is more than one or two adherend thicknesses (essentiallythe only case of practical interest) are given by
B
cos y
0, B(l + 3k) cash [@(x - L ) / t ] I4 sinh @A/) 3 mu(' k,
x-lo-l
+ ucos y,
The maximum stresses in the bond layer are given by: Maximum bond shear stress,
zb)-
=
0;
+ mU(l - k)/tanh Maximum bond peel stress,
lo+l-x + sin y, t
1
(28.26)
($)I
(28.27)
Adhesive joints 643 Figure 28.41 gives a comparison of the maximum bond stresses as functions of the loading stress ax for two different adherend thicknesses in a joint with a bond layer thickness of 0.01. It is interesting to note that the peel and shear stresses take on quite similar values. Since the maximum peel stress varies approximately as y i according to eqn (28.27) (the contribution of U being relatively minor), the relationship for ys given in eqn (28.26) suggests that the peel stresses should be expected while the same variation is to vary as (t/tb)1/2, seen from eqn (28.27) for the maximum shear stresses since p also contains (t/t,)'/* as a factor. Thus both stresses should vary with the thickness ratio by the same factor. The fact that they are numerically close together for all stresses is partly due to the effect of other parameters that enter into eqns (28.27) and (28.28) and partly due to the fact that k does not vary much with load for axgreater than 5. A slight nonlinearity can be observed in the curves of Fig. 28.41 for the lower loading stresses. Figure 28.42 gives a comparison of maximum bond stresses in single and double lap joints for a fixed value of the loading stress ax. For loading stresses above this value the bond
stresses vary essentially in proportion to the load even in the single lap joint, as just discussed. The stresses are plotted in this figure as a function of adherend thickness with the adherend axial modulus as a parameter. The trend toward higher bond stresses and therefore a greater tendency toward bond failure with increasing adherend thickness which was discussed in Section 28.4.2 is clearly borne out in these curves. Note also that reduction of the adherend modulus tends to aggravate the bond stresses. In addition it is apparent that there is considerable separation between the peel and shear stresses in the case of the double lap joint, the peel stresses for the latter case being smaller. This reflects the fact that the peel stresses vary linearly as yd defined in eqn (28.18b) and therefore vary as (t/tb)1/4 rather than as (t/tb)l/*as in the single lap case. Thus, peel stresses for double lap joints are not as much of a factor in joint failure as they are in single lap joints, although they are still large enough relative to the shear stresses that they can not be ignored. Failure characteristics of single and double lap joints will be discussed below. If the adherends are thin enough, failure in double lap joints should be in the form of adherend
&=looOO, Gb=150,%=SO0 '4
g=10
M.02; (52,
I
12
4
/
lo
E " si3
0
10
20
-
=*
30
40
Fig. 28.41 Maximum bond stresses in single lap joint, bond thickness = 0.01.
644 Mechanical fastening and adhesive bonding Double Lap Joint
Single Lap Joint
peel strau
4 -
3.5 -.
Shear Stress
Ex = 10,OOO Ex = 5,wO
3-
2.5
-
-------
2-
' 1.5 -
O -0
1
0.5
0'5 0
0.02
0.08
0.04
0.1
0
0.02
0.06
AdherendThickness
Ex I20,oOo
0.04
0.06
0.08
AdherendThickness
+
0.1
CTx= 10 Gb= 150 E,= 500
l,,= 50 1 = 10
Fig. 28.42 Maximum bond stresses in single and double lap joints, fixed vx = 10.
axial (tensile or compressive) failure. For single lap joints, adherend bending stresses are significant at the ends of the overlap as indicated in Fig. 28.39; using standard beam formulas, the maximum axial stress for combined bending and stretching (the latter stress corresponding to the single lap joint in tension loading) for the bending deflection given in eqn (28.22) can be expressed as QJrnax = OX3(1+ t,/t)k
(28.29)
The maximum adherend axial stress is largest for adherends which are particularly thin with respect to the bond thickness; these will be prone to brittle bending failures for composite adherends or to yielding associated with bending for metal adherends. Hart-Smith discusses difficulties with the use of standard single lap shear test specimens50.The problem is that adherent bending failures are likely to occur with such specimens rather than bond failures and test results obtained in such cases tend to be irrelevant and misleading. One additional characteristic difference between single and double lap joints should be discussed. The effect of lateral deflections
on single lap joint performance is quite long range. Figure 28.43 shows that for a joint with an adherend thickness of 2.54 mm (0.1 in) the bond stresses do not reduce to their minimum values until the overlap length reaches a value in the range of 10-12.7 cm (4-5 in), for a loading stress of 69 MPa (10 ksi). Double lap joints also require some minimum length before stresses settle out as a function of overlap length, but in this case the stresses reach minimum values with respect to overlap length for lengths on the order of 5 to 10 adherend thicknesses, in the present case amounting to 1.3-2.5 cm (0.5-1 in).
28.4.3.4.2 Effects of adherend tapering In this section we will consider joints with adherend thicknesses which vary along the joint length. These include the configurations shown in Figs. 28.44 and 28.45, namely, double strap joints with tapered outer adherends and scarf joints as well as step lap joints. As discussed in Section 28.4.2.2, tapering the outer adherends of strap joints as in Fig. 28.44(a) is beneficial mainly for reducing or eliminating
Adhesive joints 645
-
t m e w 0.7
m
1.6 1
.'
0.5
..
Signmba--lO Gb--160 Eb--500 Er--S.OOO tb- 0.01 M call~ ht@h
-
-- 20
Fig. 28.43 Effects of overlap length in single lap joints.
(A) PARTIALLY TAPERED STRAP JOINT
Qf-
I am i nat e
I
meta 1
I Triangular
Elewnt I
Fig. 28.45 Generic step lap joint.
+B *
I
(E) SCARF JOIN1 I
Bond
Fig. 28.44 Tapered double strap and scarf joints.
peel stresses, while scarf and step lap joints (Figs. 28.44@), 28.45) can eliminate shear stress peaks as well as peel stresses. With both tapered outer adherends and scarf joints, it can be shown that the bond stresses can be related to the ratio of taper length to thickness by
zb= axt/Z ;ab=axt 2 / P
(28.30)
This relationship is quite accurate for scarf joints having the same maximum stiffness
(axial modulus times maximum thickness) in each adherend. For tapered portions of strap joints it is fairly accurate for the peel stresses, and holds approximately for the shear stresses if the tapered portion is not too long. Note that eqn (28.30) implies that the bond stresses are constant along the length of the joint and can be reduced to any arbitrary level by making t/Z small enough, i.e. making the joint long enough with respect to the adherend thickness. Moreover, the effect of t/Z on the peel stresses is quite strong, being governed by the square of the thickness-to-length ratio. This is especially important in the case of outer adherend tapering in strap and lap joints as a means of reducing peel stresses to a manageable level. Step lap joints (Fig. 28.45) represent a compromise version of the scarf joint which can
646 Mechanical fastening and adhesive bonding take advantage of the layered structure of the composite adherend. The average slope of the region represented by the line through the steps in Fig. 28.45 tends to control the average shear stresses developed in the bond. Within each horizontal section, equivalent to the tread of a staircase, the behavior is analogous to a joint with constant adherend thickness, and the differential equation given in eqn (28.10) (Section 28.4.3.2)applies locally when tu and tL are adjusted to match the local situation. An expression similar to eqn (28.12), i.e. for thejth step,
+-
PB
gives the maximum shear stresses within each step, and the overall solution is a chain of such expressions with allowance for continuity of the shear strain and resultants, Tuj and TLjat the points where neighboring steps join. In each step of the joint the shear stresses will have a distribution similar to that of Fig. 28.36, the size of the peaks being governed primarily by the length of the step through the paramecan ter P.Z./t.The aspect ratio for the step, Zj/t, 11 in prmciple be kept small enough to almost completely avoid any peaking by using a large number of steps and keeping the length of each one small. In practice, the number of steps is governed by the number of plies in the laminate. In addition, if the joint is used to connect a composite adherend to a metal component, machining tolerance requirements and cost considerations for the metal part enter into the selection of the number of steps. The following discussion will address the specific benefits of adherend tapering in
tapered double strap joints, scarf joints and step lap joints. The overall approach is to aim for a highly efficient joint which reduces the effects of shear and peel stress concentrations at the ends of the joint. Ideally we would like to achieve the joint strength provided by the 'P over A' concept obtained with the case discussed in Section 28.4.3.2 for perfectly rigid adherends (Fig. 28.27), in which increasing the joint length indefinitely brings the shear stress in the bond down to any required level regardless of the magnitude of load being supported by the joint. While tapering does reduce the peel stresses markedly in the tapered strap joint as will be seen below, shear stress peaks can not be avoided, and the law of diminishing returns continues to prevail with regard to increasing the joint length to obtain greater load capacity; however, adhesive ductility will enhance the strength beyond what elastic analysis suggests. Double strap joints with tapered outer adherends are considered in Figs. 28.46 to 28.48. Figure 28.46 indicates the tapered configurations that are considered. Figure 28.47 gives some shear stress predictions for joints with uniform adherend thickness for comparison with the tapered cases which are considered in Fig. 28.48. Figure 28.46 defines the notation used in Fig. 28.45 in terms of 'fully tapered' outer adherends (Fig. 28.46(a)), partially tapered adherends in which the taper extends only part of the length of the joint (Fig. 28.46@))and fully tapered adherends with an 'initial rise', i.e. in which the thin end of the adherend does not come to zero thickness. (The term '% initial rise' implies that the rise is expressed as a percentage of the maximum adherend thickness.) The three cases considered in Fig. 28.48 can be compared with the case for uniform adherends with equal upper and lower adherends modulus (E, = EL)in Fig. 28.47. For the situation of no initial rise, two cases are considered in Fig 28.48, the case of 50% taper and that of full taper. There is an appreciable difference in the shear stress distribution at the
Adhesive joints 647 left end of the joint for these two cases, but the peel stress distribution is essentially unaffected. For both the full taper and 50% taper cases, a minor tensile secondary peel stress peak is present at the right end of the region under consideration (near the midpoint of the strap joint). The peel stress expression in eqn (28.30) gives a good estimate of the peel stress level at the left end of the joint, and the result of the estimate is so small for both cases that the difference is not distinguishable in Fig. 28.48. However, in the case of an initial rise of only 1 / 4 of the maximum outer adherend thickness (25% initial rise), significant peel stresses arise at the left end of the joint, in fact, about 80% of the level occurring for the case of no tapering. The initial rise also causes a greater increase in shear stress at the left end of the joint than in the case of 50% taper. Thus, tapering is advantageous mainly as a way of eliminating the effects of peel stresses in double strap joints. Once this is accomplished, the effects of peel stress peaks can be controlled to a significant extent by taking advantage of adhesive ductility. Tapered strap joints can not achieve the ideal behavior which is possible with scarf or step lap joints, but they provide a simpler solution to good joint performance if the adherends are thin enough. Shear stress distributions in scarf joints (Fig. 28.49) are given in Fig. 28.50. Practical scarf
(A) F u l l y tapered - - no i n i t i a l r i s e
(@ 5G% tapered j o i n t ---I
1.0
I.c I
----
I
(C] F u l l taper
/ 0.05
-
-
25% i n i t i a l r i s e
initial rise
\'T
-==E==I
Fig. 28.46 Tapered strap joints under consideration.
eT
Shear strws
- - Uniform Adhemd l h l c k m u :I
:I
0
0.2
0.4
b.r
0.8
1 X
1.2
1.4
1d
1.8
2
Fig. 28.47 Shear stresses in untapered strap joints.
648 Mechanical fastening and adhesive bonding
5 T
(A) Shear stresb
- - Double Strap Tapered Adherends
Fig. 28.48 Stresses in tapered double strap joints.
joints are arranged in a symmetric double lap configuration which avoids bending effects. Figure 28.49 represents a balanced stiffness design for dissimilar materials, by achieving a continuous thickness change over the length of the joint. The most important parameter for
-
-
--Es
/
0.2, TI, E.16
ffii
0.4 QI graphite epoxy E.8 msi
Fig. 28.49 Stiffness-balanced scarf joint configura-
tion.
the scarf joint is the effect of adherend stiffness unbalance ( E , # Ei; ’0’ and ‘i’ refer to the outer and inner adherends as in Fig. 28.36). The results given in Fig. 28.50 represent the effect of varying degrees of stiffness unbalance. The ratio of peak-to-average shear stresses compare well with the values given by Hart-Smith37,although the latter did not give the distribution of stresses along the length of the joint. For fairly sizeable unbalances, up to 4:1, the maximum shear stress peak is not as great as that observed in Fig. 28.47 for the uniform adherend case. However, it is clear that a stiffness unbalance will increase the shear stress peak and weaken the joint. For the equal stiffness case the shear stress is constant and equal to the average stress at all points.
Adhesive joints 649
EL I EU = 4
EU = 8,000 tu = n = .2 G b = l 6 0 Eb-600 t b = O . O l
-
Sigma x = 10
2
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
X
Fig. 28.50 Shear stresses in scarf joints.
If there is no reason why the joint cannot be configured as in Fig. 28.49 for dissimilar materials so as to take advantage of the benefits of the balanced stiffness case, then in principle, the scarf joint provides a near ideal solution to achieving as much load capacity as is required in any situation without overstressing the bond layer. However, the dimensions of the joint may grow too large to be practical for high joint load. In addition, an extremely good fit, for example, to tolerances on the order of the bond thickness over large lengths, has to be maintained to ensure that the joint can maintain uniform load capacity over its length. Thermal stresses will also be a factor in various combinations of dissimilar materials which will prevent the ideal form of behavior from being achieved. In terms of the HartSmith approach to avoiding creep failure under slow cyclic loading, the balanced scarf joint is at a disadvantage in not providing a shear stress minimum. For this situation the allowed load would have to be limited to prevent environmental conditions corresponding to hot wet exposures; thus the advantage of the scarf joint in eliminating stress peaks
might be lost if performance in hot wet environments is required. Step lap joints798oare treated in Figs. 28.51 to 28.53. Figure 28.51 shows a generic joint configuration that is introduced to illustrate some of the effects of design parameters on stresses in the joint. The results presented in Figs. 28.52 and 28.53 were generated for this discussion using a linear elastic response model for the adhesive; in practice, considerable strength capability of the adhesive is unused if elastic response of the adhesive is assumed; Fig. 28.54 taken from the discussion by Hart-Smithsois an example of joint design using elastic-plastic response for the adhesive. The elastic adhesive model used in Figs. 28.52 and 28.53 is adequate for illustrate some of the controlling parameters on joint design. These results are based on the classical Volkersentype analysis with provision for resultant and shear strain continuity at the interfaces between neighboring steps, as discussed previously. The 5-step design in Fig. 28.52 and the 10-step design in Fig. 28.53 were chosen with the following characteristics:
650 Mechanical fastening and adhesive bonding
I
I
/
V
/
Quasi-Isotropic Carbon E =8nsi X
Except for the first and last steps, the adherend thickness was equal for each step. The first and last thickness increments were half those of the generic steps. The lengths of each step were chosen with-a fixed value of the parameter qs, = pirj/fj, where lj is the length of step j . The half-incremented end steps gave a more uniform shear stress distribution than maintaining the same increment for all steps. The symmetric joint configuration shown in Figs. 28.51 and 28.54, the thickness increment for the outer adherend (composite) was greater than that for the inner adherend by the inverse of the modulus ratio, to achieve the desired stiffness balance for the dissimilar adherends. The parameter ETA listed in Figs. 28.52 and 28.53 is equivalent to qs, defined above. This parameter essentially controls the length of the joint; both Figs. 28.52 and 28.53 show an increase in joint length with qsl ('ETA' in the two figures). Note further that the load capacity of the joint in terms of the allowed resultant, 5, listed as "BAR in Figs. 28.52 and 28.53, which provides for the required bond shear stress limitation of 5 (ksi for the units mentioned earlier) shows a general increase with joint length, but with diminishing increase when vsl gets much beyond 3. Table 28.1 gives a summary of the results shown in the two figures.
@OXY,
I
Fig. 28.51 Step lap joint configuration.
As discussed above, the joint design shown in Fig. 28.5437represents a practical joint design which accounts for several considerations that the simplified elastic analysis approach used for Figs. 28.52 and 28.53 neglects. The neglect of ductility effects has already been mentioned. In addition, the use of as large a
Fig. 28.52 Shear stresses in 5-stepjoint.
Adhesive joints 651 presented by Corvelli and Saleme79;this was later enhanced by Hart-%~ith~~, under NASA funding, to provide for elastic-plastic response, culminating in the A4EG and A4EI programs-, to allow for variations in thickness, porosity, flaw content and moisture content in the bond layer. Hart-SmithsoI8lnotes that in mathematical treatments of step joints, all properties have to be constant within each step; however, in an actual joint such as that shown in Fig. 28.54, artificial breaks may be inserted to permit changes in porosity of bond thickness. 28.4.4 MECHANICAL PERFORMANCE OF
ADHESIVES
28.4.4.1 Ductile response of adhesives
C L : : : : : : : : : : : : : : : : 0
uyaa1mkL11unuuur I
Fig. 28.53 Shear stresses in 10-step joint.
number of steps as 10 in Fig. 28.53 may not be practical. The joint design shown in Fig. 28.54 represents the evolution of steplap joint design over many years. Early analytical work was
Figure 28.55 taken from the 1983 edition of the DoD/NASA Advanced Composites Design Guides1 show shear stress-strain response characteristics of typical aerostructural adhesives. Figure 28.55(a) represents a relatively ductile film adhesive, FM73, under various environmental conditions, while Fig. 28.55@) represents a more brittle adhesive (FM400) under the same conditions. Similar curves can be found in other sources62. Temperature dependence and strain rate dependence of the stress-strain characteristics are important characteristics; these are also addressed in the results shown in Ref. 62. Even for the less ductile material such as that represented in Fig. 28.55@),ductility has a pronounced influence on mechanical response of bonded joints, and design only for elastic response deprives the
Table 28.1 Summary of step lap joint results (Figs. 28.52 and 28.53)
No. of steps
10
10
10
5
5
1
2
3
3
6
(in)
4.44 (1.75)
8.89 (3.5)
13.33 (5.25)
6.05 (2.47)
12.5 (4.93)
Allowed resultant, kN/cm (103 lb/in)
12.03 (6.87)
18.78 (10.72)
22.05 (12.59)
12.35 (7.05)
13.43 (7.67)
V7sl
Joint length, cm
652 Mechanical fastening and adhesive bonding
7 6
5 4
3 2 1
- 0
1
3
2
5
4
0
6I
1
2
3
1 .z
f.4
4
6
Fig. 28.54 Practical step lap joint designs0.
( A ) FM78 NON P O R O U S
0
0.2
0.4
0.8
1
0.8
S h e s r 8 traln
LT = -55 d q C RT = Room Temp H T M = 60 degC1100%RH
6o 46
$ -aa
I:2
5
I$ F m o o ~ a m u r b m d
16 10
6 0
o
om
om
001
om
ai
a12
0.14
Mr-
Fig. 28.55 Typical stress-strain characteristics of aerospace adhesivess1.
6
Adhesive joints 653 application of a significant amount of avail- zp, which can be obtained by solving for zp able structural capability. from eqn (28.32), leading to The work of H a r t - S ~ n i t h emphasized ~~~ the 'p = Gb$max - d[(GbOYmax)2 - 2GboSEl importance of ductile adhesive response and introduced the relationship between the strain (28.33) energy to failure of the adhesive and the load capacity of the joint. As a means of simplifying Hart-Smith has also used an equivalent the stress analysis of the joint in the presence bilinear representation in which the horizontal of ductile adhesive response, Hart-Smith part of the curve is set equal to zmax,the maxishowed that any bilinear stress-strain curve mum shear stress of the actual stress-strain which has the same ultimate shear strain and curve, and the initial modulus Gboadjusted to maximum strain energy as that of the actual give the strain energy match, using the expresstress-strain curve will produce the same total sion: load in the joint. Figure 28.5681gives an examGbo = 'ma:('maxYmax /2 (28'34) ple of the method for fitting a bilinear curve to the actual stress-strain curve of the adhesive which is also obtained from eqn (28.32) when in shear. With the strain energy of the adhesive rmax is substituted for zp. In either case the use given by of a bilinear representation of the stress-strain SE = rJmax- zp2/2Gb0 (28.32) curve for the response of the adhesive in shear makes it straightforward to obtain one dimenwhere GbO! yma, and SE are the initial modulus sional stress distributions in various types of of the stress-strain curve, the maximum strain joint geometry with adhesive ductility and the strain energy of the adhesive at ymaX accounted for; solutions are given for single respectively, then the equivalent bilinear curve and double lap joints with uniform and consists of an initial straight line of slope Gbo tapered adherends, as well as more sophistitogether with a horizontal part at an abscissa cated joint designs such as scarf and step lap
Fig. 28.56 Elastic-perfectly plastic adhesive response model (Fh473)*'.
654 Mechanicalfastening and adhesive bonding
geometries ( H a r t - S ~ n i t h ~These ~ ~ ) . have subsequently been incorporated in the 'A4Ex' series of computer programsmentioned previously. Figure 28.57 (see notation of Fig. 28.36) shows an example of the use of the bilinear stressstrain curve approximation, in this case for predicting the stresses in a symmetrical
double lap joint with equal adherend stiffnesses (i.e. Eo = Ei; to = fi/2). Figure 28.57(a) gives the distribution of upper adherend axial stress resultant while Fig. 28.57@) gives the shear stress distribution in the bond layer. The linear portions at the ends of the resultant distribution in Fig. 28.57(a)correspond to the ends of the shear stress distribution in Fig. 28.57@)
(A) Upper adherend resultant distribution
4.5
f 4
( 1 k-lblln
= l.762 kNlom)
Axial load 310 MPa (46 ksl)
0 0
0.6
1.I
1
2
K
le
(6)Shear stress distribution (1 ksi= 6.896 MPa)
Axial k a d 310 MPa (45 ksi)
0.5
--
o i 0
I
0.c
z
1.Q
Fig. 28.57 Stress distributions in double lap joint - ductile adhesive response.
2
Adhesive joints 655 where the shear stress is a constant because of the plateau in the bilinear representation of the stress-strain curve, in agreement with the equilibrium relation given in eqn (28.6). Following the analysis developed by Hart-Smith, the lengths of the plastic zones designated in Fig. 28.57(b) as Ip are given by
lp = (aX/2zp- l/pbd)to;p,,
= [2Gbflto/EotbI’/*
then eqns (28.36a, b) give the same value. The factor (2GbOymax/zp - 1)1/2in eqn (28.36b)acts as a load enhancement factor and represents the increase of joint load capacity due to ductile adhesive response over the maximum load allowed by elastic response of the adhesive. Note that eqn (28.36b) can be rearranged to express 5x)ma, in terms of the maximum strain energy of the adhesive:
(28.35) Here p,, (subscript ’bd’ denoting balanced double lap) is equivalent to p given in eqn (28.14) when the latter is specified for the case of equal-stiffness adherends, while 5, is the nominal loading stress at either end of the overlap. The expression for Ip given in eqn (28.33)is valid only if greater than 0, of course, negative values of plastic zone length not having any meaning. Thus if /33,/2 < zp, no plastic zone is present and the behavior of the joint can be considered to be purely elastic. The maximum value of 5, for this case can be expressed by inverting the shear stress expression in eqn (28.17) with oth= 0, for the case of equal adherend stiffnesses and setting zb)max to zp. For the case of pax/2 2 zp which corresponds to ductile response of the adhesive (the plateau of the bilinear stress-strain curve), the Hart-Smith analysis35provides the required expression for 0, The two cases are summarized as follows: pbdSX/2 < zp (elastic response): ‘x)max
= 2zp/pbd
E
PbdCx/22 zp (ductile response):
where ye = zp/Gb,
(28.37)
then eqn (28.3613)can be written: 2 =
p’(2Gb()SE)
(28.38)
The Hart-Smith analysis based on the equivalent bilinear stress-strain law was shown in Ref. 35 to give the same joint load capacity as the solution for the problem using the actual stress-strain curve of the adhesive. The convenience of the bilinear stress-strain description is in the simplicity of the solutions it allows; once the length of the plastic zone at each end is determined, the same types of solution apply for the elastic zone as were given in eqns (28.9) and (28.10) for the resultant and shear stress distributions, together with linear resultant and constant shear stress distributions in the plastic zones. The most obvious effect of ductility in the adhesive behavior is the reduction of peak shear stresses. In addition, there is a beneficial effect on reduction of peel stresses. For the (28.36a) double lap joint considered in Fig. 28.57, the maximum peel stresses denoted by ob),ax which occur at the ends of the joint, are given39 by: Ob)max
(28.36b)
(
=
y = 3-
If y, = zp/GbO,which is the maximum strain in the elastic part of the bilinear representation,
;
Fb)max
i’”
;
E, = peel modulus of adhesive
(28.39)
656 Mechanical fastening and adhesive bonding where z ~ is )the ~maximum ~ ~ shear stress, leads to high shear and peel stresses at the either /?OX/2 for the elastic case or zp for the ends of the joint, and may inhibit desirable case of ductile response. The maximum peel flow characteristics of the adhesive. On the stresses are thus reduced by the same ratio as other hand, thick bond lines tend to generate the maximum shear stresses in the case of duc- porosity which weakens the bond. Data presented in Ref. 83 show a fairly persistent tile response of the adhesive. tendency for lap shear strength to drop off Even though ductile response of the adhesomewhat as the bond thickness is increased sive provides additional load capacity of the above 0.12 mm (0.005 in). joint over what is provided by purely elastic In addition to effects of bond thickness per response, it is advisable to keep the load se, Hart-Smith60 discusses the effect of bond capacity of the joint low enough to ensure thinning at the ends of the joint which is purely elastic response for most practical situcaused by resin flow during curing. Figure ations where time-varying loading is 28.58 illustrates the tendency toward bond encountered. Some damage to the adhesive thinning at joint edges together with some probably occurs in the ductile regime which manufacturing techniques for avoiding the sitwould degrade the long-term response. The uation. Loss of bond thickness may cause main benefit of ductile behavior is to provide considerable elevation of shear and peel increased capacity for peak loads and damage stresses in the bond. In addition to the tolerance with regard to flaws - voids, porosapproaches shown in Fig. 28.58, tapering of ity and the like - in the adhesive layer. In the adherends near the ends will help to alleaddition, calculations of the plastic zone length play a part in the avoidance of creep viate the situation; tapering from the inside failures which can constitute a major consider- surface of the adherend will also provide a ation for slow cyclic loading in hot wet local thickening of the bond line to compensate for thinning due to resin loss. environments. Effects of porosity in the bond layer are illustrated in Fig. 28.59(j1which compares the response of FM73 for porous and non porous 28.4.4.2 Effects of bond layer defects bond layers for various environmental condiDefects in adhesive joints include surface tions. The data6' indicate that porosity is preparation deficiencies leading to low mainly a characteristic of thickness of the bond strength interfaces between the adhesive and layer. There is some loss of structural capabiladherends, voids and porosity, and lack of ity in the presence of porosity in the bond, but bond thickness control. Surface preparation there may still be adequate strength for the effects were discussed in Section 28.4.2.6 and bond to function as required if the joint is will be treated in considerable detail in designed adequately. Since porosity is associChapter 29. However, it should be kept in ated with thickened regions of the bond which mind that adhesive joints will not succeed in tend to occur away from the edges, porosity providing dependable performance if good tends to be confined to the interior of the joint surface preparation procedures are not main- where the stresses are relatively small, and may not be objectionable in many cases. The tained. Bond layer thicknesses of 0.12-0.25 mm main focus in Ref. 61 is the effect of adherend (0.005-0.010 in) are typical of structural thickness with regard to damage tolerance in bonded joints. There appears to be a tradeoff the presence of bond layer defects. The issue between negative effects which occur when has to do with the design principle discussed the bond is too thin and those occurring for in Section 28.4.2.1 of keeping adherend thicktoo thick bond layers. Thinness in bond layers nesses within limits which ensure that the
Adhesive joints 657
n V
4. NIIOIQFfCERVVACVUWMQ
Fig. 28.58 Manufacturing techniques to relieve bond pinch-off 51,
60
I rLT
/Stress VQ. Strain
50
2 I 40
-- FM731
Y
67 67
E5 5
- -55dqC
30
LT -
20
RT -room temp.
-
H/w
- - 60 degC/lOo% rh
10
2.
o $ 0
I
0.5
1
1.5
2
S h e a r Strain
Fig. 28.59 Effect of porosity on adhesive stress-strain characteristicss1x : porous; : non porous bond layers.
adherends fail rather than the bond. Ductile response of the adhesive has an important influence on the situation. By making use of eqn (28.36b) above, together with the definition of p,, given in eqn (28.35), the adherend thickness limit for a double lap joint,
tJc,which ensures adherend failure can be expressed by restating eqn (8) of Ref. 52 as follows:
-
658 Mechanical fastening and adhesive bonding damage, while tests conducted at one cycle per hour produced failures within a few hundred t, = 4 {Maxadhesive strain energy]/ cycles. On the other hand, specimens repre{Maxadherend (elastic) strain energy] (28.41) sentative of structural joints which have the Hart-Smith states that for to not greater than characteristic shear stress trough seen in Figs. to),, given in eqn (28.40),joint performance will 28.37 and 28.57 are able to sustain hot wet conbe relatively insensitive to bond flaws pro- ditions even at low cycling rates if the length vided there is adequate ductility in the of the elastic region (Ie in Fig. 28.57) is long adhesive response. Problems may arise with enough. Based on experience of the PABST high temperature adhesives such as the FM program, the Hart-Smith criterion for avoid400 considered in Fig. 28.55(b), since these ance of creep failure requires that t,),, be no tend to have limited ductility. As indicated in greater than t p / l O . But the stress analysis for eqn (28.36b), limited ultimate strain capability the elastic-plastic case using the bilinear adhe(y,,,) will reduce margin of ultimate strength sive response model leads to an expression for the minimum shear stress equivalent to eqn over elastic response of the joint. (28.12)with 2 replaced by Ze:
which is equivalent to:
Two major considerations in the joint design philosophy of Hart-Smith are: (1)either limiting the adherend thickness or making use of more sophisticated joint configurations, such as scarf and step lap joints, to ensure that adherend failure takes precedence over bond failure; (2) designing to minimize peel stresses, either by keeping the adherends excessively thin or, for intermediate adherend thicknesses, by tapering the adherend. In addition, it is essential that good surface treatment practices be maintained to ensure that the bond between the adhesive and adherends does not fail. When these conditions are met, reliable performance of the joint can be expected for the most part, except for environmental extremes, i.e. hot wet conditions. The HartSmith approach focuses primarily on creep failure associated with slow cyclic loading (i.e. one cycle in several minutes to an hour) under hot wet conditions, this corresponds, for example, to cyclic pressurization of aircraft fuselages. In the PABST program4143,18 test specimens used for characterizing adhesives (so-called 'thick adherend' specimens) which are designed to produce essentially uniform shear stress along the bond were tested at high cycling rates (30 Hz) and were able to sustain more than 10 million loading cycles without
' P sinh PbdZe/2fo , I -
28.4.4.3 Durability of bonded joints
*b)&
(28.42)
do,, - see eqn (28.35)).Since sinh (3) = 10, this amounts to a requirement that PbdZe/2tobe at least 3, i.e. that the elastic zone length be determined by Ze 2 6to/Pbd.Since le is equivalent to the total overlap length, 1, minus the sum of the plastic zone lengths, i.e. 21, then P making use of the expression for ZP m eqn (28.33), the criterion for elastic zone length reduces to a criterion for total overlap length corresponding to a lower bound on 1 which can be stated as
(
-
1 2 -+-to
4 pbd
)
(28.43)
Equation 28.43 for the joint overlap length is the heart of the Hart-Smith approach to durability of bonded joints for cases where adherend failure is ensured over bond failure for static loading and in which peel stresses are eliminated from the joint design. This type of requirement has been used in several contexts. For example60,it becomes part of the requirement for acceptable void volume in the bond layer, since in t h s case the voids, acting essentially as gaps in the bond layer, reduce the effective length of the overlap. The criterion has to be modified numerically for joints other than symmetric double lap joints with equal
Adhesive joints 659 stiffness adherends and uniform thickness. For more sophisticatedjoint configurations such as step lap joints, the A4EI computer code provides for a step length requirement equivalent to that of eqn (28.43) for simple double lap joints. In addition to creep failures under hot-wet conditions, the joint may fail due to cracking in the bond layer. Johnson and Mall7*presented the data in Fig. 28.60 which shows the effect of adherend taper angle on development of cracks at ends of test specimens consisting of composite plates with bonded composite doublers, at lo6 cycles of fatigue loading; here the open symbols represent the highest load levels at which cracks fail to appear while the solid symbols are for slightly higher loads at which cracks just begin to appear. It is noted that even for outer adherend taper angles as low as 10” (left-most experimental points in Fig. 28.60) for which peel stresses are essentially nonexistent for static loading, crack initiation was observed when the alternating load was raised to a sufficient level. A number of factors need to be clarified before the implications of these results are clear. In particular it is of interest to establish the occurrence of bond cracking at shorter cycling times, say less than 3 x lo5 cycles corresponding to expected lifetimes of aircraft. Effects of cycling rate and environmental exposure are also of interest. Nevertheless, the data presented in Ref. 65
FM-XII
MBOND
D NO MBOND APPLIED CYCLIC
S I A E S S , IM ~ 5 MPa ,
0
PREDlClED
,,.* -M TAPER ANGLE.
60 0.
Po
deg
Fig. 28.60 Crack development in bonds of tapered composite doublers at lo6loading cycles7*.
suggest the need for consideration of crack growth phenomena in bonded composite joints. Indeed, a major part of the technical effort that has been conducted on the subject of durability of adhesive joints6”” has been based on the application of fracture mechanics based concepts. The issue of whether or not a fracture mechanics approach is valid needs further examination.Apparently, no crack-like failures occurred in the PABST program, which was a metal bonding program, even when brittle adhesives were examined at low temperatures. The amount of effort which has been expended by a number of respected workers on development of energy release rate calculations for bonded joints certainly suggests that there is some justification for that approach, and the results obtained by Johnson and Mall appear to substantiate their need for composite joints in particular. 28.4.5 MECHANICAL BEHAVIOR OF COMPOSITE ADHERENDS
28.4.5.1 Joint failure characteristics
Typical failure modes in structural joints are illustrated in Fig. 28.61 which are indicative of adherend rather than bond failures. In the case of single lap joints (Fig. 28.61(a))bending failures of the adherends will occur because of high moments at the ends of the overlap. For metal adherends, bending failures take the form of plastic bending and hinge formation, while for composite adherends the bending failures are brittle in nature. In the case of double lap joints, peel stresses build up for thicker adherends causing the types of interlaminar failures in the adherends illustrated in Fig. 28.61(b). 28.4.5.2 Thermal stress effects
Thermal stresses are a concern in joints with adherends having dissimilar thermal expansion coefficients. Figure 28.62 illustrates the
660 Mechanical fastening and adhesive bonding
c
A.
A 0. AND C INDICATE FAUURh I - E
(B) Double Lap hints
Fig. 28.61 Failure modes in composite adherends49, 50.
E; = E, 68.97GPa (10000 ksi); G, = 1.04 GPa (150 ksi); f1/2 = f, = 2.54 mm (0.1 in); fb = 0.254 mm (0.01 in); a, = 23.4 x lo4 "C-'(13 x 1O4 OF-'); a. = 1.8x 1O4"C (1 x 10-6°F-1 ). Cure temperature 121.1"C (250°F); Application temperature 23.9% (75°F) Loading stress 68.97MPa (10 ksi).
Fig. 28.62 Thermal shear stresses in double lap joints: outer adherend 0 / 9 0 carbon epoxy; inner adherend aluminum.
Adhesive joints 661 effect of thermal stresses in a double lap joint consisting of an aluminum inner adherend and a 0/90° carbon epoxy outer adherend. The stresses due to thermal mismatch between the aluminum and composite arise if the cure temperature of the bond is substantially different from the temperature at which the joint is used. The case considered here represents a 121°C (250°F) cure temperature for the adhesive and a room temperature application, a temperature difference of -79°C (-175"F), which (see Tables 28.2 and 28.3) would result in a strain difference of 0.002 between the aluminum and composite if no bond were present. (The material combination considered here, aluminum and carbon epoxy, represents the greatest extreme in terms of thermal mismatch between materials normally encountered in joints in composite structures.) Thermal stresses in bond layers of double lap joints can be determined from the expressions given in eqns (28.14-28.20). (These calculations are all based on an assumed elastic response of the adhesive.) Hart-Smith3"39 provides corrections for ductile response in the presence of thermal effects. Figure 28.62 illustrates how the thermal stresses combine with the stresses due to structural load to determine the actual stress distribution in the adhesive. The thermal stresses in themselves develop an appreciable fraction of the ultimate stress in the adhesive, and although they oppose the stresses due to structural loading at the left end, they add at the right end and give a total shear stress that is somewhat beyond the yield stress of typical adhesives, even with as small a structural
loading stress as 69 MPa (10 ksi). Similar effects occur with the peel stresses, although the peel stresses due to thermal mismatch alone have the same sign at both ends of the joint; with a composite outer adherend the thermally induced peel stresses are negative, which is beneficial to joint performance. Peak peel and shear stresses obtained from these relationships for various combinations of metal and composite adherends whose properties are given in Tables 28.2 and 28.3 are shown in Table 28.4. For joints with an aluminum inner adherend, the difference in thermal expansion between the adherends is relatively large, giving considerably higher thermal stresses for the most part. In addition, carbon epoxy has a particularly low thermal expansion, which tends to produce higher thermal stresses with carbon epoxy adherends in combination with metals than do other composites. Note that boron epoxy in combination with titanium gives particularly small thermal stresses because of similarity of the thermal expansion coefficients shown in Tables 28.2 and 28.3 for these materials. As Table 28.3 Generic metal properties (MIL-HDBK-5 1983)
Ti6-Al4-4V
1025 Steel
2014 Aluminum
110.3
206.9
69.0
Poisson ratio
0.3
0.3
0.3
a, 10" OC-1
8.82
10.26
23.4
Young's modulus, GPa
Table 28.2 Generic mechanical properties of composites (C.C. Chamis NASA Lewis Research Center, NASA TM-86909,1985)
Unidirectional lamina Composite Boron epoxy S-glass epoxy Carbon epoxy
E,, GPa
E , GPa
vLT
aL,10-6 "C-'
201 60.7 137.9
20.1 24.8 6.90
0.17 0.23 0.25
11.7 3.78 0.72
0/90 Laminate
aT,1 P
OC-l
30.4 16.7 29.5
E x GPa
ax1 k6O C - I
113.8 43.72 72.6
7.92 2.34
8.6
662 Mechanical fastening and adhesive bonding Table 28.4 Bond layer thermal stress in double lap joints (0/90 composite outer adhered, metal inner
adherend) Boron epoxy
Glass epoxy
Carbon epoxy
0.419 0.061 -0.465 -0.067
2.33 0.338 -3.73 -0.541
15.64 2.27 -19.43 -2.817
5.44 0.789 -6.30 -0.914
7.99 1.16 -15.0 -2.17
26.22 3.80
27.7 4.02 -24.4 -3.54
28.2 4.08
-
Titanium Shear stress, MPa
(ksi) Peel stress, MPa (ksi) Steel
Shear stress, MPa (ksi) Peel stress, MPa (ksi)
-38.1 -5.52
Aluminium
Shear stress, MPa (ksi) Peel stress, MPa
(ksi)
40.1 -5.82
40.47 5.86 -44.6 -6.47
t, = 5.08 mm (0.2 in); to adjusted for equal adherend stiffnesses;t, = 0.253 nun (0.1 in). Adhesive properties: shear modulus, 1.03 GPa (150 ksi); peel modulus, 3.49 GPa (500 ksi). Give temperature, 121°C (250°F).Application temperature, 24°C (75°F).
discussed earlier, the 'peel' stresses shown in tion for metal adherends which are relatively Table 28.4 are all negative (i.e. compressive) stiff with respect to transverse shear deformabecause of the location of the composite on the tion, but for polymer matrix composite outside of the joint, although the shear stresses adherends which have low transverse shear are unaffected by this aspect of the joint. moduli, transverse shear deformations are Composite repair patches on aluminum air- more significant and can have an important craft structures benefit from this type of influence on bond layer shear stresses. A usebehavior, in that peel stresses are not a prob- ful correction to the classical Volkersen lem for temperatures below the cure solution which allows for transverse shear temperature. Placing the metal rather than the deformations in the adherends can be composite on the outside of a double lap joint obtained by modifying the shear modulus of would reverse the signs of the peel stresses the adhesive from its actual value, Gb, to an making them tensile and aggravating the effective value, Gb)eff,given by effects of differential thermal expansion of the (28.44) adherends. 28.4.5.3 Transverse shear and stacking sequence effects in composite adherends
where Ksh = 1 +
Classical analyses such as the Volkersen shear lag model for shear stresses in the bond layer (Sections 28.4.3.3 and 28.4.3.4.1) are based on the assumption that the only significant deformations in the adherends are axial, and that they are uniformly distributed through the adherend thicknesses. This is a good assump-
Here GUo and Gxzi are the transverse shear moduli of the adherends. For the double lap joint, the parameter /?appearing in eqn (28.14) (see the third equation of the top row of eqn (28.14)) is then modified by replacing Gb by Gb/K,,, using the value for Ksh given in eqn (28.44), and all the expressions in eqns
References
663
3. Oplinger, D.W. and Gandhi, K.R., Analytical (28.15-20) for stresses in the bond layer are studies of structural performance in mechanimodified by the resulting alteration of p. The cally fastened fiber-reinforced plates. In Proc. correction given here amounts to treating 1/ 3 A r m y Solid Mechanics Conf. 1974, Army the thickness of each adherend as an extension Materials and Mechanics Research Center of the bond layer, and assigning the shear stiffManuscript Report AMMRC MS 74-8 (1974). ness of the adherend for that part of the 4. Garbo, S.P., Ogonowski, J.M. and Reiling, H.E., effective bond layer. The factor 1/3 correJr, Effect of variances and manufacturing tolerances on the design strength and life of mechanically fassponds to a linear distribution of shear stress tened composite joints. v2 Air Force Wright through the adherend thicknesses, which is Aeronautical Laboratories Report AFWAL-TRconsistent with the assumption that the axial 81-3041 (1981). deformations are approximately uniform 5. Hyer, M.W. and Klang, E.C., Contact stresses in through the adherend thickness. pin-loaded orthotropic plates, Virginia Tech Center As an example, consider joint with a 0/90 for Composite Materials and Structures Report carbon epoxy outer adherend joined to an aluCCMS-84-02 (1984). 6. Ramkumar, R.L., Saether, E.S. and Appa, K., minum inner adherend, with adherend Strength analysis of laminated and metallic plates thicknesses of 2.53 mm (0.1 in) and 5.06 mm bolted together by many fasteners, Air Force Flight (0.2 in), respectively, and a 0.253 mm (0.01 in) Dynamics Laboratory Report AFWAL-TR-86bond thickness. Assume a shear modulus of 3034 (1986). the bond layer of 1.06 GPa (150 ksi) and trans7. Madenci, E. and Illeri, L., Analytical determinaverse shear moduli of 4.82 GPa (700 ksi) for the tion of contact stresses in mechanically fastened composite adherend and 26.5 GPa (3800 ksi) composite laminates with finite boundaries. Intern. J. Solids Sfructures 30, pp. 2469-2484 for the aluminum. A value of 1.839 is then (1993). obtained for Ksh, and the value of p and the 8. Crews, J.H. and Naik, R.A., Combined bearing maximum shear and peel stresses which and bypass loading on a graphite/epoxy lamidepend on it are reduced by a factor of (Ksh)1/2 nate. Composite Structures, 6, pp. 2148 (1986). or 1.36 for this case. The shear and peel 9. Hart-Smith, L.J., Mechanically-fastened joints stresses are therefore approximately 30% for advanced composites - phenomenological lower than the values predicted with the considerations and simple analyses. In Fibrous Composites in Structural Design. New York: unmodified bond shear modulus. This type of Plenum Press (1980)pp. 543-574. correction can be shown to give relatively 10. Petersen, R.L., Stress Concentration Factors. New good predictions of the adhesive stresses in York: John Wiley and Sons (1974)p. 135. comparison with finite element analyses. In 11. Lenoe, E., Oplinger, D.W. and Burke, J,J., addition, the departure of Ksh given in eqn Fibrous Composites in Structural Design. In (28.44) from 1 gives a good indication of the Proc. 4 f h Con& Fibrous Composites in Structural range of joint parameters for which adherend Design, New York: Plenum Press (1980). 12. Oplinger, D.W., On the Structural Behavior of shear deformations are important. REFERENCES
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664 Mechanical fastening and adhesive bonding (1944). stress concentrations. J. Composite Materials, 8, 28. Kutscha, D. and Hofer, K., Feasibility of joining pp. 253-265d (1974). advanced composite Fight vehicle structures, Air 15. Garbo, S.P. and Ogonowski, J.M., Efect of variForce Materials Laboratory Report AFML-TRances and manufacturing tolerances on the design 68-391 (1968). strength and life of mechanically fastened composite joints, Air Force Flight Dynamics Laboratory 29. Dickson, J.N., Hsu, T.N. and McSkinney, J.N., Development of an understanding of the fatigue pheReport AFFDL-TR-78- 179 (1978). nomena of bonded and bolted joints in advanced 16. Chang, EK. and Hung, Chang-Li, Response of filamentary composite materials. Vol. 1, Analysis double shear-lap bolted composite joints under mulMethods, Lockheed Georgia Aircraft Company, tiple bypass loads. Standford University Dept. of USAF Contract Report AFFDL-TR 72-64, vol I Aeronautics and Astronautics Report, Doctoral (June 1972). dissertation of Chang-Li Hung. (1993). 17. Garbo, S.P., Efjrects of bearingbypass load interac- 30. Grimes, G.C., Wah, T. et al., The development of non-linear analysis methods for bonded joints in tion on laminate strength, Air Force Flight advanced filamentary composite structures, SouthDynamics Laboratory Report AFWAL-TR-81West Research Institute, USAF Contract Report 3114 (1981). AFML-TR-72-97 (September 1972). 18. Naik, R.A. and Crews, J.H., Jr., Ply-level failure analysis of a graphitelepoxy laminate under bear- 31. Renton, W.J., The analysis and design of composite materials bonded joints under static and fatigue ing/bypass loading, NASA Technical loadings, PhD Thesis, University of Delaware Memorandum 100578 (1988). (1973). 19. Hart-Smith, L.J., Design methodology for bondedbolted composite joints. v l Air Force Wright 32. Renton, W.J. and Vinson, J.R., The analysis and design of composite materials bonded joints under Aeronautical Laboratories Report AFWAL-TRstatic and fatigue loadings, Air Force Office of 81-3154 (1982). Scientific Research Report TR-73-1627 (1973). 20. Ramkumar, R.L., Saether, E. and Cheng, D., Design guide for bolted joint in composite struc- 33. Oplinger, D.W,. Stress analysis of composite joints. In Proc. 4th Army Materials Technology tures, Air Force Wright Aeronautical Con$, Newton, MA: Brook Hill Publishing Co. Laboratories Report AFWAL-TR-88-3035 (1986). (1975), pp. 405-451. 21. Shivakumar, K.N. and Crews, J.H., Jr., Bolt clamp-up relaxation in a graphitelepoxy laminate, 34. Hart-Smith, L.J., AFFDL-TR-72-130, pp. 813-856. NASA Technical Memorandum 83268 (1982). 22. Slepetz, J.M, Oplinger, D.W. and Andrews, B. 35. Hart-Smith, L.J, Adhesive bonded double lap joints, NASA Langley Contractor Report, NASA CRO., Bolt tension relaxation in composite friction 112235 (1973). joints. In Proc. 7th Con$ Fibrous Composites in Structural Design, Wright Patterson 36. Hart-Smith, L.J., Adhesive bonded single lap joints, NASA Langley Contractor Report, NASA CRAeronautical Laboratories Report AFWAL-TR112236 (1973). 85-3094 (1985). 23. Harris, H.G., qalvo, I. and Hooson, R.E., Stress 37. Hart-Smith, L.J., Adhesive bonded scarf and stepped-lap joints, NASA Langley Contractor agd deflection analysis of mechanically fastened joints, Air Force Flight Dynamics Laboratory Report, NASA CR-112237 (1973). 38. Hart-Smith, L.J., Analysis and design of advanced Report AFFDL-TR-70-49 (1970). composite bonded joints, NASA Langley 24. Shyprykevich, P., Characterization of bolted joints Contractor Report, NASA CR-2218 (1973). behavior, MIL-HDBK-17 Accomplishments at Standardization ASTM J. Composite Technical 39. Hart-Smith, L.J., Advances in the analysis and Research, 17(3),pp. 260-270 (1995). design of adhesive-bonded joints in composite 25. Military Handbook MIL-HDBK-17: Polymer aerospace structures. In SAMPE Process Matrix Composites (1992). Engineering Series, 19, pp. 722-737 Asuza: 26. Volkersen, O., Die Nietkraftverteilung in SAMPE, (1974). Zugbeanspruchten Nietverbindungen mit 40. Primary adhesively bonded structure (PABST) technology, Air Force Contract F33615-75-C-3016 Konstanten Laschenquerschnitten. Lufffahnforschung, 15, pp. 4-47 (1938). (1975). 27. Goland, M. and Reissner, E., Stresses in 41. Thrall, E.W., Primary adhesively bonded structure cemented joints. J. Appl. Mech., 11, A17-27 technology (PABST) Phase 7 b: Preliminary Design,
References 665 Air Force Flight Dynamics Laboratory Report AFFDL-1R-76-141 (1976). 42. Shannon, R. W. et al., Primary adhesively bonded structure technology (PABST): General Material Property Data, Air Force Flight Dynamics Laboratory Report AFFDL-TR-77-101 (1977). 43. Land, K.L., Lennert, F.B. et a/., Prima ry adhesively bonded structure technology (PABST): Tooling, Fabrication and Quality Assurance Report, USAF Technical Report AFFDL-TR-79.3154 (October, 1979). 44. Hart-Smith, L.J., Adhesive bond stresses and strains at discontinuities and cracks in bonded structures. Trans. J. Engng Mater. Tech., 100, pp. 128-144 (1978). 45. Hart-Smith, L.J., Differences between adhesive behavior in test coupons and structural joints, Douglas Aircraft Company Paper 7066. Presented to ASTM Adhesives Committee D- 14 Meeting, Phoenix, Arizona, 1981. 46. Hart-Smith, L.J., Design methodology for bondedbolted composite joints, Douglas Aircraft Company, USAF Contract Report AfWAL-TR81-3154, Vol I and I1 (February 1982). 47. Thrall, E.W., Jr., Failures in adhesively bonded Structures, AGARD-NATO Lecture Series No. 102, ‘Bonded Joints and Preparation for Bonding’, Oslo, Norway and The Hague, Netherlands, April 1979 and Dayton, Ohio, October 1979. 48. Hart-Smith, L.J., Further developments in the design and analysis of adhesive-bonded structural joints, Douglas Aircraft Company Paper 6922. Presented at the ASTM Symp. Joining of Composite Materials, Minneapolis, MN, April 1980. 49. Hart-Smith, L.J., Adhesive bonding of aircraft primary structures, Douglas Aircraft Company Paper 6979. Presented to SAE Aerospace Congress and Exposition, Los Angeles, California, October 1980. 50. Hart-Smith L.J., Stress analysis: a continuum analysis approach. In Developments in Adhesives - 2 (ed. A. J. Kinloch), London: Applied Science Publishers, pp. 1-44(1981). 51. Hart-Smith, L.J. and Bunin, B.L., Selection of taper angles for doublers, splices and thickness transition in fibrous composite structures. In Proc. 6th Conf. Fibrous Composites in Structural Design, Army Materials and Mechanics Research Center Manuscript Report AMMRC MS 83-8 (1983). 52. Nelson, W.D., Bunin, B.L. and Hart-Smith, L.J.,
Critical joints in large composite aircraft structure. In Proc. 6th Conf. Fibrous Composites in Structural Design, Army Materials and Mechanics Research Center Manuscript Report AMMRC MS 83-8 (1983). 53. Oplinger, D.W., A layered beam theory for single lap joints, US Army Materials Technology Laboratory Report MTL TR 91-23 (1991). 54. Oplinger, D.W., Effects of adherend deflections on single lap joints. Int. J. Solids Structures, 31 (18), pp. 2565-2587 (1994). 55. Hart-Smith, L.J., Adhesively bonded joints in fibrous composite structures, Douglas Aircraft Paper 7740. Presented to the Intern. Symp. Joining and Repair of Fibre-Reinforced Plastics, Imperial College, London (1986). 56. Hart-Smith, L.J., Induced peel stresses in adhesivebonded joints, Douglas Aircraft Company, Technical Report MDC-J9422A, August 1982. (see also USAF Report AFWAL TR-82-4172, (1982).) 57. Hart-Smith, L.J., Brown, D. and Wong, S., Surface preparations for ensuring that the glue will stick in bonded composite structures, 10th DoD/NASA/FAA Conf. Fibrous Composites in Structural Design, Hilton Head Is, SC (1993). 58. Hart-Smith, L.J., Ochsner, W. and Radeckv, R. L., Surface preparation of fibrous composites for adhesive bonding or painting. Douglas Service Magazine, 1, pp. 12-22 (first quarter 1984). 59. Hart-Smith, L.J., Ochsner, W. and Radecky, R. L., Surface preparation of fibrous composites for adhesive bonding or painting. Canadair Service News, 2, pp. 2-8 (1985). 60. Hart-Smith, L.J., Effects of adhesive layer edge thickness on strength of adhesive-bonded joints, Quarterly Progress Report No. 3, Air Force Contract F33615-80-C-5092 (1981). 61. Hart-Smith, L.J., Effects of flaws and porosity on strength of adhesive-bonded joints, Quarterly Progress Report No. 5, Air Force Contract F33615-80-C-SO92 (1981). 62. Frazier, T.B. and Lajoie, A.D., Durability of adhesive joints, Air Force Materials Laboratory Report AFML TR-74-26, Bell Helicopter Company (1974). 63. Becker, E.B. et al., Viscoelastic stress analysis including moisture difision for adhesively bonded joints, Air Force Materials Laboratory Report AFWAL-TR-84-4057 (1984). 64. Jurf, R. and Vinson, J., Efects of moisture on the static and viscoelastic shear properties of adhesive
666 Mechanical fastening and adhesive bonding joints, Dept. of Mechanical and Aerospace Engineering Report MAE TR 257, University of Delaware (1984). 65. Mostovoy, S., Ripling, E.J. and Bersch, C.F., Fracture toughness of adhesive joints. J. Adhesion, 3, pp. 125-144. (1971). 66. DeVries, K.L., Williams, M.L. and Chang, M.D., Adhesive fracture of a lap shear joint. Experimental Mechanics, 14, pp. 89-97 (1966). 67. Trantina, G.G., Fracture mechanics approach to adhesive joints, University of Illinois Dept. of Theoretical and Applied Mechanics Report T&AM 350, Contract N00019-71-0323 (1971). 68. Trantina, G.G., Combined mode crack extension in adhesive joints, University of Illinois Dept. of Theoretical and Applied Mechanics Report T&AM 350, Contract N00019-71-C-0323 (1971). 69. Keer, L.M., Stress analysis of bond layers. Trans. ASME J. Appl. Mech. E., 41, pp. 79-83 (1974). 70. Knauss, J.F., Fatigue life prediction of bonded primary joints, NASA Contractor Report NASACR-159049 (1979). 71. Wang, S.S. and Yau, J.F., Analysis of interface cracks in adhesively bonded lap shear joints, NASA Contractor Report NASA-CR- 165438 (1981). 72. Johnson, W.S. and Mall, S., A fracture mechanics approach for designing adhesively bonded joints. In Delaminafion and Debonding of Materials, ASTM Special Technical Publication STP 876, American Society for Testing and Materials, pp. 189-199 (1985).
73. Tsai, M.Y. and Morton, J., On numerical and analytical solutions to the single lap joint. Intern. J. Solids and Structures (1994). 74. Benson, N.K., Influence of stress distribution on strength of bonded joints. In Adhesion, Fundamentals and Practice, New York: Gordon and Breach, (1969), pp. 191-205. 75. Adams, R.D. and Wake, W.C., Structural Adhesive Joints in Engineering, Amsterdam: Elsevier Applied Science Publishers (1984). 76. Kuenzi, E. and Stevens, G., Determination of mechanical properties of adhesives for use in the design of bonded joints, Forest Products Laboratory Note FPL-011 (1963). 77. Snedon, I., The distribution of stress in adhesive joints. In Adhesion, (ed. D.D. Eeley), Ch. 9, Oxford: Oxford University Press (1962). 78. Carpenter, W., Goland and Reissner were correct. J. Strain Analysis, 24(3), pp. 185-187 (1989). 79. Hart-Smith, L.J., Further developments in the design and analysis of adhesive-bonded structural joints. In Joining of Composite Materials, American Society for Testing of Materials Special Technical Publication ASTM STP 749 (1981). 80. Hart-Smith, L.J., In Fiber Composite Analysis and Design, Federal Aviation Administration Technical Center Report DOT/FAA/CT-88/18, Vol. 2, Ch. 3 (1988). 81. DoD/NASA Advanced Composites Design Guide, 1983.
SURFACE PREPARATIONS FOR ENSURING THAT THE GLUE WILL STICK IN BONDED COMPOSITE STRUCTURES
29
L.J. Hart-Smith, D. Brown and S. Wong
29.1 INTRODUCTION
29.2 HISTORICAL BACKGROUND
Adhesively bonded joints can be no stronger than the interface between the adhesive and the members being bonded together. In-service bond failures have always been associated with weak interfaces, for both metal and composite adherends. While most people acknowledge that adhesive bonding of metallic structures requires strict adherence to proper processes, many people unthinkingly accept the notion that it is easy to make epoxy stick to epoxy, for example, and pay no attention to the need for proper processing for the adhesive bonding of composite structures. This chapter begins with a historical review of the need for appropriate surface treatment to ensure that the glue will stick to composite surfaces. It then focuses on photomicrographs of different surfaces, to which the adhesive will or will not stick, as a basis for inspections prior to bonding. Such prebond inspections are vital because of the inability to detect weak bonds after manufacture, until they have fallen apart. The characterization of the surfaces prepared by different techniques can assist in formulating process specifications that will ensure reliable adhesive bonding and in identifying past practices that should be discontinued.
During the late 1970s and early 1980s, the Douglas Aircraft Company at Long Beach, California, was the site of one of the technically most successful research contracts ever funded by the Wright-Patterson Air Force Laboratories - the Primary Adhesively Bonded Structure Technology (PABST) program. This research was directed at well-known problems concerning the adhesive bonding of metallic aircraft structures. These problems were the need to change contemporary processing from etching to anodizing, and the need to select adhesives and primers on the basis of long-term durability rather than short-term strength. These failings had been made very clear by widespread in-service problems experienced for many years by both commercial and military operators. A successful outcome to the research was assured because of the success of the Redux bonding developed in England during WWII and since employed extensively by Fokker. The PABST program succeeded in all of its objectives, bar one. It omitted a large-scale flight demonstration program, because it would have solved no problem that had not already been solved by the successful ground testing of panels and a complete wide-body fuselage barrel with simulated wing center section. However, in
Handbook of Composites.Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
668 Suface preparationsfor ensuring that the glue will stick in bonded composite structures retrospect, such a demonstration was necessary as a public relations exercise to convince the decision makers to apply the technology, where appropriate, on the primary structure of large aircraft. Nevertheless, the technology did not die. The combination of phosphoric acid anodizing for aluminum alloys, a phenolic-based corrosion-inhibiting primer and adhesives formulated to resist the absorption of water has since been applied very extensively on the primary structures (wings, fuselages, and tails) of Cessna and SAAB aircraft with excellent results. The composites bonding industry could be described today as being in the same state that metal bonding was in 25 years ago, but with one big difference. The need to properly prepare the surfaces of composite laminates prior to bonding is acknowledged only by those who have suffered a major problem from having failed to do so on some previous occasion. Ironically, there are many researchers and production engineers worldwide who are aware of the problems and their causes. However, there is no highly visible activity like the PABST program to draw attention to the issues. This chapter cannot be expected to solve these problems, but it is hoped that it will raise the level of awareness of the subject and that its content will help achieve better bonded composite structures by providing standards for inspecting the surfaces before bonding, rather than creating the impossible situation of trying to detect weak bonds after bonding. What is needed is a method that will ensure a reliable bond every time, using procedures to which all manufacturers of composite structures will be willing to adhere. The first part of this challenge has already been accomplished: light grit-blasting or thorough mechanical abrasion has been shown to work. Unfortunately, more often than not, films of adhesive are applied to the surfaces left by the removal of a peel ply without any surface treatment. Most, and possibly all, peel plies are known not to create a suitable surface for bonding, at least when used with laminat-
ing resins or adhesives. They create a weak interface, with failure occurring at only a fraction of the strength that would have been developed by grit-blasting. However, the resulting weak bonds cannot be detected by conventional ultrasonic nondestructive inspection. Also, most bonded composite structures are so thin that they will not fall apart even if the bond has as little as one-tenth of the strength of a properly processed bond. Consequently, little has been done about the problem, in the mistaken belief that doing so would incur an unnecessary expense. On the contrary, even if local bond failures can be tolerated structurally, once they become detectable in service they cause expensive inspection programs and possibly rework, not necessarily to add to the strength of the parts, but to prevent the flaws from being detected and repaired again on subsequent inspections. Failing to ensure that the glue sticks properly in the first place is definitely a false economy. The delay in solving the corresponding problems on bonded metal structure gave the technology such a bad name in the USA that it was the direct cause of the extensive use of riveted structure when bonding would have been lighter, less expensive, and more durable, with better than a twentyfold reduction in the incidence of fatigue cracks and a dramatic improvement in damage tolerance. Lest the same preference for mechanical fastening rather than bonding continue to be followed for thin composite structures, it should be noted that the typical fasteners cost about a hundred times as much as the simple rivets used with aluminum structures. Worse, the minimum diameter of the specialty fasteners for composites is 50% larger than the diameter of comparable rivets. Conventional aluminum rivets are not used for carbon composite structures because of concern about galvanic corrosion, between the rivets and the carbon fibers, so composite structures that could have been bonded must be locally reinforced to be mechanically fastened instead, adding to both the cost and weight.
Historical background 669 There are, then, very great incentives for reliable processing of adhesive bonds in composite structures. There is also no hstory of properly processed composite bonds failing in service. (This is also true of metal bonding. All in-service failures are the result of environmental attack, at load levels far too low to have initiated mechanical failure of the bonds. Testing during the PABST program of coupons cut from retired aircraft stored at DavisMonthan, Arizona, showed that there was no structural deterioration after 20 years.) There are others that call for improvements in surface preparation for bonding of composites. Robert Schliekelmann, the famous pioneer of Redux bonding at Fokker, was sufficiently concerned about the failure to recognize the need for proper surface treatment of composites as well as metals to make a plea’ for more attention to this issue. Almost a decade ago, Douglas engineers prepared an article on the subject2to help the airlines until the repair manuals could be updated. The article was judged to be of sufficient importance to be reprinted in the Canadair house journal3.Every experiment had to be repeated to create new photographs and every phenomenon was duplicated, thereby verifying that the problems discussed in the original article really existed. A similar concern was expressed in England, where Parker and Waghom4 reported on a far more comprehensive test program on the effects of surface preparation on adhesive bond strength for carbon-epoxy laminates. They also concluded that ‘composite surfaces must be abraded to achieve strong adhesive-bonded joints.’ Pocius and Wentz advocated the use of Scotchbrite pads with embedded abrasive particles as an effective and reliable technique for achieving good composite bonds5. A recent research program6 addressed both thermoset and thermoplastic composites. Reference 6 also cites other English research. The problems still exist, on both sides of the Atlantic Ocean. A more recent article on this subject7,voiced concerns because of the reintroduction of inferior techniques throughout much of the
aerospace industry. It seems significant that, today, some factories prohibit the use of bonding directly to a composite surface created by the mere removal of a peel ply and that the automotive racing industry has experienced the same kind of premature failures with peelply treatments that the aerospace industry has suffered. If what may be called the peel-ply bonded joints were twice as strong as they actually are, there would be little concern. Conversely, if they were only half as strong, the problem would have been more widely recognized long ago and far more done about it. The real problem with peel-ply bonded joints is that, while quite unreliable in the eyes of those closest to the subject, they usually have sufficient strength to pass ultrasonic inspections (because there are no gaps) and are sufficiently strong some of the time that the joints do not fail prior to delivery of the component. This makes it difficult to present a convincing case to non-technically minded participants in the industry that there is a real problem. One non-failure tends to be interpreted as complete success. This chapter cannot possibly provide all the answers on the subject of surface preparations for composite bonding. It took a $20 million 5 year program to solve the corresponding problems for metal bonding. It should also be acknowledged that the concerns expressed here are based exclusively on consistently bad experiences with 180°C (350°F) cured epoxy composites. (The problem may be associated with the curing of the laminate and the peel ply as much as with the adhesive.) It may also be significant that most, if not all, of these problems were associated with nylon peel plies, whose use with a 180°C (350°F) curedlaminate is believed by some polymer chemists to be undesirable, because the nylon may break down and transfer a thin film of low-molecular weight material to the composite surface. Photographic evidence of this phenomenon is included here. Such a surface would be very difficult for an epoxy adhesive to wet, because of too low a surface energy.
670 Surface preparationsfor ensuring that the glue will stick in bonded composite structures The polymer chemists express less concern about the use of nylon at 120°C (250"F), but some adhesion failures of glued joints have occurred with those materials also. The dominant factor may be the low surface energy level of the composite cured against a slick peel ply, or one contaminated by a release agent to ensure that it can be removed easily without fracturing (damaging) the composite laminate. It is certain that the criteria by which manufacturers evaluate their peel plies are precisely the opposite of those that someone else trying to promote adhesion in a bonded joint would follow. Some peel plies are even coated with release agents to ensure that they can be removed easily without damaging the basic laminate. Silicone transfer has been observed with some peel plies supposedly totally free of contamination. What is needed to create a bondable surface is a tear ply that will remove a small amount of surface resin over the entire area to be bonded. However, peel plies known to be capable of achieving this are almost invariably shunned because they are so hard to peel off, because they break and inevitably lead to local contamination as the process is restarted, or because there is concern that the interior of the laminate will be damaged. It cannot be asserted on purely technical grounds that there are no circumstances under which some peel plies will produce an adequate surface for adhesive bonding. It stands to reason that some users of this approach have been spared the problems Douglas and others have encountered, or the products would have all been removed from the marketplace. However, it is likely that still others have also suffered, but are unaware of the cause of their problems. Nevertheless, the argument against peel plies is irrefutable on economic grounds. The cost of investigating weak bonds has been far, far greater than would have been incurred if Douglas had insisted that all suppliers and subcontractors lightly grit-blast or thoroughly hand-sand composite surfaces following removal of the
peel plies. Not one in-service delamination has occurred in those composite structures prepared for bonding by the grit-blast technique. It is hoped that this chapter will inspire those who believe that their structural adhesive bonds are both strong and reliable to assess composite surfaces prior to bonding. (A request for information from a major US aerospace manufacturer about a peel ply different from those used at Douglas resulted in their switching to another peel ply. Microscopic examination of the surface to be bonded made it clear to them that one of their problems could be tied directly to the choice of a new peel ply. Once alerted, they checked all related procedures and even changed one of their old and trusted peel plies once its true capabilities had been exposed.) This chapter concentrates exclusively on the issue of surface preparation. Obviously, one must also follow correct thermal and pressure profiles during cure to achieve a properly cross-linked adhesive that has a chance to flow and wet the surface to which it is to be bonded. Prebond moisture, in both laminates and adhesives, is also known to be a major cause of weak bonds. These and other important issues are discussed in References 7 and 8. This does not imply that they are any less important. We now present electron-micrographs of glue that failed to stick to the composite, composite surfaces to which the adhesive did not stick, similar surfaces to which the glue will not stick, surfaces to which adhesive is known to have stuck in the past, and surfaces to which it is hoped it will stick. The effectiveness of light grit-blasting as a reliable surface treatment has already been established. It was used on all bonded joints of the Lear Fan all-composite aircraft9, creating bonds stronger than the surrounding structure even when half the overlapping area was disbanded because of inappropriate solid-rubber tooling used for the fuselage skin splices. The bond to the gritblasted surface was so strong, where the mating surfaces were brought properly into
The problem - weak bonds 671 contact, that 100%bonding was not necessary. Grit-blasting is used today on some of the most trouble-free composite components on the MD-11 aircraft; specifically, the components made in Japan. However, some manufacturers of bonded composite structures will prefer to continue to use peel-ply-only surface preparations, no matter how weak or unreliable the resulting bonds are. One must find different peel plies that do not release cleanly, but which take some of the surface resin with them, without damaging the remainder of the matrix, or find a coupling agent to improve the behavior of what are known today to be unsatisfactory peel plies no matter how widely they may be used, or transfer the production to factories in which reliable surface-treatment practices are followed.
area with a different texture is part of the surface ply of unidirectional carbon-epoxy left when the remainder separated. The reason why this very small segment adhered is unknown. It is almost as if there were one small drop of chemical acting like a wetting agent on the composite when it was bonded. If this should prove to be the case, and research can identify an agent capable of promoting the wettability of cured epoxy in a laminate by uncured epoxy in an adhesive layer, it would be a giant contribution to composites technology. The idea of a coupling agent, equivalent to the silanes used for epoxy-bonding of aluminum alloys during repairs, is apparently feasible for composites also - at least when the peel ply has not been coated with a release agent. Coupling agents would be more likely to work if the basic laminate were not fully cured prior to bonding. 29.3 THE PROBLEM - WEAK BONDS One would prefer an incomplete initial cure in No more convincing proof of the existence of order to leave some active chains in the moleweak bonds created on peel-ply composite cular structure of the composite to which the surfaces can be found than in Fig. 29.1. This is glue could bond. The bond cycle would then a scanning electron microscope (SEM) pho- be relied upon to complete the cure of the tomicrograph not of a composite surface after laminate and this process might affect the the peel ply was lifted off, but of the cured choice of adhesive. However, such an outadhesive, showing the imprint left by the com- come must be preceded by an posite laminate after it peeled off! The small acknowledgment that the remainder of Fig. 29.1 and the large similar surrounding area imply the existence of a problem. Figure 29.2 shows the same lack of adhesion evident in Fig. 29.1 at a small magnification over a very much larger area, to show how widespread this phenomenon can be. The texture of the peel ply is clearly imprinted in the glue over almost the entire area. The different texture visible on the left side of the figure is the peel-ply imprint on the underlying composite laminate. In this area, the adhesive failed to bond to the lower surface. Throughout everywhere else shown in Fig. 29.2, the adhesive failed to bond to the other composite part, the peel-ply imprint of which is embossed on the adhesive. Figure 29.3 Fig. 29.1 Imprint of peel ply, in background, on shows an adjacent area on the same-part. The cured adhesive film that failed to adhere (mag x30). thick glue layer (shiny area) at the bottom left
672 Surface preparations for ensuring that the g2ue will stick in bonded composite structures
Fig. 29.2 Large area of adhesive that failed to bond, showing imprint of peel-ply surface (mag X5).
of Fig. 29.3 shows the unmistakable imprint of the fine-weave peel ply left by the other composite part where it did not stick. Immediately above it, there is a small area of cohesive fracture of the bond. (The reason for bonding in just that small area is not known.) The dark area in the lower right of Fig. 29.3 shows the imprint of the peel ply in the composite underneath the adhesive layer. The small vertical white lines define the fine weave of the peel ply. The more widely spaced larger white blotches, with some sharply defined edges, reveal the very small extent of the sanding carried out on this sample. The white blotches are the high points in the coarser weave of the plain-weave cloth in the basic laminate. These are evident on the surface opposite from that cured against the lay-up tool. These large white blotches in Fig. 29.3 illustrate one unavoidable problem with sanding as a surface preparation. How can one sand down to the bottom of every depression in the fine-weave texture left by the peel ply without sanding into the fibers at the top of every raised bump in the texture of the composite material itself? This problem is aggravated by the use of woven fabrics rather than unidirectional tape and is unavoidable on the bag side of the laminate, unless a caul plate is used on both sides. The customary use of dry (nonimpregnated) peel plies, which absorb resin from the surface plies of the laminate, results in such an ultrathin layer of surface resin that it is
Fig. 29.3 Adhesive that failed to bond, also showing sanding confined to high spots of the composite weave (mag ~ 5 ) .
virtually impossible to sand the surface without breaking the fibers. One needs to use preimpregnated peel plies or to include a layer of resin (or adhesive) film between the basic lay-up and any dry peel ply to create what amounts to a gel coat sufficiently thick to sand as the surface for bonding or painting. It is also impossible to completely sand a surface containing ply drop-offs or local buildups. Only light grit-blasting would work in such a case. It should be noted that, at the time when the bond shown in Figs. 29.1 to 29.3 was manufactured, conventional ultrasonic inspection failed to find any evidence of a defect. The gap had not yet opened up even though an afterthe-fact inspection revealed almost no adhesion throughout most of the area meant to have been bonded. This is very significant. Weak bonds cannot be detected by inspection before they have failed. They can be prevented only by adhering to process specifications that ensure they will not occur. Figure 29.4 shows the typical composite surface left by removal of the same type of peel ply (in this case, a corona-treated nylon peel ply). A comparison with Fig. 29.1 will leave no doubt about the origin of the surface embossed into the cured adhesive layer. (Scientific testing of this particular piece of adhesive revealed nothing untoward about the way it was cured, although there are
The problem - weak bonds 673
Fig. 29.4 Imprint of nylon peel ply in epoxy surface of composite laminate (mag x65).
unconfirmable suspicions that prebond moisture may have exacerbated the problem.) An examination of the surface of the nylon peel ply itself after removal from the laminate revealed no resin adhering to the peel ply (Fig. 29.5). [Regrettably, this is not from the same sample, even though it is nominally the same peel ply cured in the same laminating resin, at 180°C (350"F).]Significantly, the other side on the peel ply was covered by a considerable amount of resin bled from the surface plies of the laminate. This is most undesirable, because most composite specifications call for unreasonable resin starvation as the norm. (The purpose of minimizing the resin content is to boost the easily measured in-plane tension strength. Resin-dominated strengths are both more difficult to measure and more difficult to analyze, so few such considerations are included in the design process.) There is usually barely enough resin to hold the fibers together. Removing surface resin makes it more difficult to transfer loads through bonded joints. It also makes the surface more prone to impact damage and more difficult to prepare for repainting in service. This is acknowledged by the way in which some Boeing components are now made with preimpregnated peel plies despite the added cost and storage and handling problems with respect to dry ones.
Fig. 29.5 Enlargement of nylon peel ply after stripping off composite laminate cured at 180°C (350°F) (mag X41).
Higher resin contents in aerospace composite structures do not necessitate no-bleed cure cycles. They retain a higher resin content, but they also prevent the removal of any volatiles absorbed and trapped before cure or generated during the cure. The lack of volatile removal introduces a whole new family of processing problems and structurally inferior laminates. Figure 29.6 shows a greatly magnified view of the same peel ply shown in Fig. 29.5. The peeling motion was from right to left. (The corresponding picture to the left of the vertical bundle of peel-ply fibers shows far less fracturing of the matrix.) Noteworthy features shown in Fig. 29.6 include (1) the tendrils of resin around the vertical filament on the extreme left, (2) the way the resin has pulled away cleanly from the horizontal filaments on the right of the adjacent vertical fiber, (3) striations in the fracture surface of the resin between the second and third horizontal filaments from the top, (4)the loose piece of resin between the first
674 Surface preparationsfor ensuring that the glue will stick in bonded composite structures in thick test coupons with impractically short bonded overlaps, provided that the surface treatment and cure are carried out properly. The resulting fracture surface is shown greatly magnified in Fig. 29.7. The randomly oriented fibers are the carrier in the adhesive layer and quite distinct from the regular patterns of woven composite laminate and peel plies. There is no possibility of mistaking this cohesive fracture surface for either the interfacial or interlaminar failures shown in Fig. 29.1. I Apart from the grossly dissimilar surface textures, most adhesives are distinctively col! ored. Carbon-epoxy is invariably black. It is very easy to distinguish between:
Fig. 29.6 Highly magnified view of peel ply removed from carbon-epoxy laminate cured at 180°C (350°F) (mag x41).
and second horizontal filaments from the bottom, and (5) the white markings on the nylon filaments themselves. At still higher magnification, these markings appear to be crazing that occurs within the nylon filaments as they were bent while the peel ply was being stripped off. The marks do not appear to be resin extracted from the surface of the laminate. There was no indication of any matching roughness in the grooves left in the laminate for the samples that did match this particular piece of peel ply, although, as discussed later, there is some indication at very high magnification that ultrathin layers of nylon transferred to some areas (but not all) of the composite laminate as it cured. (There was also no evidence of nylon transfer when nominally the same peel ply was used with a different resin matrix and cured at 120°C (250°F). Had the adhesive in Fig. 29.1 stuck properly, failure would have occurred interlaminarly within the thin carbon-epoxy composite laminate, with no adhesive visible anywhere. One can enforce a cohesive failure of adhesive layers
1. Cohesive failure of the adhesive with complete coverage of both surfaces by the adhesive layer. 2. Adhesion failure with all of the adhesive on one surface, although the adhesive may fracture and end up as small segments adhering to both surfaces, but only to one in any given area (Fig. 29.8). The figure shows predominantly the surface of the adhesive to which the overlaying piece of composite
Fig. 29.7 Cohesive fracture within adhesive bond (mag x83).
Samples of diferently prepared surfaces for bonding 675 3. Interlaminar failure of the composite laminate, with fibers on both surfaces. (Failure close to the surface of the composite may leave a transparent layer of resin matrix covering the adhesive, but microscopic examination will reveal a very different fracture surface from that associated with the adhesion failure shown in Fig. 29.1.)
Naturally, there can be no photograph equivalent to Fig. 29.1 for a grit-blasted surface because the interface will not fail unless it was deliberately contaminated prior to bonding. 29.4 SAMPLES OF DIFFERENTLY PREPARED SURFACES FOR BONDING
Fig. 29.8 Predominantly adhesion failures in peelply bonded joint (mag ~ 5 ) .
failed to adhere and local dark exposed areas of the underlaying composite with the peel-ply imprint made clearly visible as those portions of the adhesive lifted off with the upper piece of composite.
The remainder of this chapter will provide a comparison between photomicrographs of different surface preparations, good and bad. Figure 29.9 shows the surface left by light grit-blasting with alumina grit at 140 kPa (20 psi). Although resin has been removed from the entire surface, so little has been removed that the texture of the peel ply is still evident. The uniformity of the surface treatment, prepared under less than ideal conditions with no handling aids, is impressive and suggests that the procedure is not unreasonably demanding, particularly when done with suitable equipment. A very highly magnified photo (Fig. 29.10) confirms that no damage was done to the fibers in the carbon-epoxy laminate. The grooves in this figure are not carbon fibers; they are the furrows left by removing the peel ply. The same grit-blast machine used to produce the sample illustrated in Figs. 29.9 and 29.10 had many years earlier burnt holes in 6.3mm (0.25 in)-thick laminates in only 20 s when operated at 698 kPa (100 psi). Restricting the blast pressure is critical when using this technique. Because the negatives of the photos used in the earlier article2 had been lost, the blast pressure was turned up to provide a comparative illustration of what happens when the blasting is overdone. Despite these
676 Surface preparations for ensuring that the glue will stick in bonded composite structures
Fig. 29.9 Lightly grit blasted composite surface, retaining imprint of peel ply (mag X50).
Fig. 29.10 Highly magnified grit-blasted epoxy surface, showing no damage to underlying fibers (mag x1000).
explicit instructions, the process did not seriously damage the laminate. The new overblasted surface is shown in Figs. 29.11 and 29.12 at low and high magnifications. The greatest difference with respect to Figs. 29.9 and 29.11 is that the texture visible in Fig. 29.11 is that of the weave of the carbon fabric. The imprint of the peel ply has been totally removed. (Figure 29.11 is at a far lower magnification than Fig. 29.9.) The vertical bands in Fig. 29.12 are the fibers. Only a tiny fraction of the surface fibers is damaged and the surface is rough enough for the adhesive to stick. The difference in size between the weaves of peel plies and typical composite fabrics is made very clear in the peculiar failure surfaces shown in Fig. 29.13. The uppermost areas are of the imprint of the fine-weave peel ply, to which the adhesive refused to bond. The lower coarser weave shows exposed carbon fibers left after very local interlaminar failures in the uppermost ply. The reason for such rapidly alternating failure modes is unknown, although it is probably significantly related to the direction of propagation of the disbond
Fig. 29.11 Deliberately over-blasted epoxy composite surface, showing weave of carbon fibers with no trace of the fine-weave peel-ply surface (mag x9).
since the exposed carbon fibers are all oriented in the same direction.
Samples of diflerently prepared surfaces for bonding 677 The lightly done and deliberately overdone grit-blasted surfaces in Figs. 29.9 to 29.12 are consistent with the ease and reliability with which the grit-blasting was carried out on the Lear Fan all-composite aircraft. A portable low-pressure machine (Fig. 29.14) was used then. It had a vacuum collector to minimize the spread of debris, and the central nozzle was surrounded by a bristle comb, which ensured that the nozzle was always held at the same distance from the surfaces. The samples shown in Figs. 29.9 to 29.12 were made under far more difficult circumstances - with cumbersome gloves to hold the part and the nozzle inside a chamber. Production parts as good as those in Figs. 29.9 and 29.10 should be far easier to produce. Figures 29.15 to 29.17 show low, intermediate, and high magnifications of the surface left Fig. 29.12 Highly magnified over-blasted compos- by removal of a nylon peel ply from a carite surface still showing minimal damage to fibers bon-epoxy laminate cured at 180°C (350°F) (mag X500). The imprint is consistent with the intermediate magnification shown in Fig. 29.4. The imprint left by each fiber in the peel ply is so smooth and slick that the adhesive will not
Fig. 29.13 Disbonded surface, showing peel-ply imprint where adhesive failed to bond and local interlaminar failures in underlying composite (mag X15).
Fig. 29.14 Portable low-pressure grit-blasting machine.
678 Surface preparationsfor ensuring that the glue will stick in bonded composite structures
I
I Fig. 29.15 Peel-ply imprint on epoxy surface to which the glue will not adhere (mag X38).
Fig. 29.17 Same sample at higher magnification (mag x750).
adhere there. It can adhere only to the exceedingly narrow strips of fractured resin between each fiber depression. The adhesive can adhere to the entire composite surface shown in Fig. 29.9, a bondable area that is an order of magnitude greater, at least. Figure 29.17 shows something not evident in the companion photo of the peel ply surface itself (Fig. 29.6). At a magnification of ~750, some of the furrows appear to be coated in places by ultrathin coatings of nylon from the peel ply. The lesser magnification of x215 (Fig. 29.16) of the same area on the surface indicates that some but not all other areas are similarly coated. Figure 29.18 shows what some regard as adequate sanding. Yet, on close examination, it is evident that the peel-ply imprint is clearly visible except for a miniscule fraction of the total area. The sanded area lies outside the area meant to be bonded, on the left, but there is no reason to suspect any different degree of Fig. 29.16 Magmfied peel-ply imprints showing transfer of ultra-thin layer of nylon to otherwise sanding. Unfortunately, many specifications call for light (or scuff) sanding because of smooth furrows in composite surface (mag x225).
Samples of diflerently prepared surfaces for bonding 679
Fig. 29.18 Inadequate abrasion achieved by scuff sanding (mag ~ 1 2 ) .
stiffener as in the unbonded area to the right. Figure 29.19 shows slight damage to the uppermost fibers as the result of thorough sanding of another sample, this time a unidirectional carbon-poxy laminate, prepared at a different facility. In this case, sanding parallel to the surface fibers would minimize any damage. However, that option would not be available for woven fabrics because the surface fibers do not all run in the same direction. The number of fibers damaged in Fig. 29.19 is an extremely small fraction of even the surface ply. Carbon fibers are so small that each tow contains 3000-12 000 individual fibers in a layer 0.13 mm-0.33 mm (0.005-0.013 in) thick. The loss of strength from damaging a few surface fibers through too energetic sanding is far less than from a bonded structure coming apart in service because of inadequate sanding. Figures 29.20 and 29.21 show sanded Samples. The woven texture is that of the peel ply, not the carbon fabric. The matt sanded areas in the figure are the high spots in the laminate surface. The peel-ply imprints still visible are in depressions. Clearly, the operator never came close to damaging the carbon fibers. Sanding should have continued for far longer or a more abrasive grade of emery paper should have been used. It is clear that some 30% of the bond area was sanded, a huge
greater concern that the fibers in the composite laminate not be damaged than for the need for sufficient mechanical abrasion to ensure that the adhesive will stick properly. As long as fifteen years ago, the McDonnell (St. Louis) specifications explicitly called for sufficient sanding to eliminate the imprint of the peel ply. In other words, the entire surface of the composite laminate was to be stripped back, just as in Fig. 29.9. (One wonders what motivated such a precise specification. Something taught them a lesson they planned not to forget.) The results of the inadequate level of sanding are evident in Fig. 29.18. The imprint Fig. 29.19 Slight fiber damage caused by thorough of the peel ply is just as clear to the left of the sanding of unidirectional carbon-epoxy laminate broken glue fillet at the edge of the disbonded (mag X460).
680 Surface preparations for ensuring that the glue will stick in bonded composite structures
Fig. 29.20 Moderately sanded peel-ply imprint on surface of woven composite laminate, showing how sanding does not abrade the entire surface (mag
Fig. 29.21 Highly magnified surface, showing how most of the peel-ply imprint remains after hand sanding (mag ~1000).
x100).
improvement over the few percent removed in the sample shown in Fig. 29.18. While it must be acknowledged that both grit-blasting and hand-sanding can be overdone, doing so takes time and effort if one is using the right abrasives and equipment. A significant loss of strength from such actions is far less likely than from either simply removing a peel ply or sanding the composite surfaces far too lightly. To cover the possibility that the bad experiences with peel plies at Douglas might have been associated exclusively with the breakdown of nylon at too high a curing temperature, samples were obtained from Oxford Brookes University, which used a different laminating resin and both a polyester peel ply and what is probably the same nylon peel ply. The polyester peel ply was noticeably more difficult to remove than the nylon peel ply, but far easier than what were referred to as tear plies in Reference 2. The surface created Fig. 29.22 Imprint of polyester peel-ply, showing by removing the polyester peel ply from a more fractured resin than with nylon peel-plies 120°C (250°F) cured carbon-poxy laminate (mag ~ 5 0 ) .
Samples of diferently prepared surfaces for bonding 681 (Fig. 29.22 mag x37.5) is generally very similar to the surface shown in Fig. 29.4 after a nylon peel ply had been removed, except for the randomly oriented tendrils, which appear to be lengths of attached polyester filaments. The university’s nylon peel ply left slick imprints almost exactly like the one used at Douglas. A comparison of the SEM photographs suggests that the polyester peel ply performed better because the filaments were smaller and there was proportionately a greater length of the very fine strips of fractured resin between the filaments. The improvement in strength from this source alone should be about 25%. However, there was another possible difference: the higher magnification of the polyester peel-ply imprint (Fig. 29.23) shows faint streaks in the grooves left by removal of the peel ply. If these could be traced to surface roughness in the resin, rather than to smooth irregularities in the polyester filaments, it would be an indication that the peel ply stripped off some surface resin and promoted adhesion of the glue.
Fig. 29.23 Highly magnified imprint of polyester peel-ply in epoxy composite (mag X500).
Fig. 29.24 Highly magnified image of polyester peel-ply after removal from 120°C (250°F) cured carbon-epoxy laminate (mag X830).
Unfortunately, a microscopic examination of the removed peel ply itself suggests that the imprint remains smooth at the molecular level. The polyester peel ply seems to be better than the nylon one but not in the same class as the surface created by light grit-blasting. Figure 29.24 shows a x830 enlargement of the surface of the removed polyester peel ply. The fibers are generally smooth and, where they are not, the visible strands represent tearing of the fibers rather than a buildup on the surface as the result of tearing the matrix resin. Nevertheless, it should be noted that the tearing shown could not have occurred if there were no adhesion at all between the peel ply and the resin matrix. The interface definitely seems more bondable than the much smoother one associated with nylon peel plies. What may be a far more sigruficant difference is that the surface created from peeling off a nylon peel ply from the carbon-poxy laminate cured at 120°C (250°F) shows no sign of nylon on the surface at either low or high magnification (Figs. 29.25 and 29.26). Nevertheless,
682 Surface preparationsfor ensuring that the glue will stick in bonded composite structures the surfaces are so slick that one would have no confidence that a reliable bond could be made other than to the thin strips of fractured resin between the grooves left by the peel ply. However, it might well be far less weak than would result from trying to bond to a nylon interface and, indeed, lap-shear testing of such bonded joints has been reported in Reference 6 as resulting in cohesive failures at the expected bond strength. The fractured surfaces look excessively porous, but so do the companion coupons made with grit-blasted surfaces, which developed essentially the same strength. These bonds definitely do not look as strong as that shown in Fig. 29.7, but they are ever so much stronger than that shown in Fig. 29.1. Figure 29.27 shows the fractured surface of an interlaminar failure within the carbon-epoxy composite. The fibers shown Fig. 29.25 Slick surface left by removal of nylon are carbon and are noteworthy for the failure peel-ply from 120°C (250°F) cured carbon-epoxy of the resin to adhere to them over most of laminate (mag X46). their surface. The fractured surface is that of the resin matrix. One would expect that the same kind of surface would be created by the 1
I -1 1
---
Fig. 29.26 Highly magnified image of nylon peel-ply imprint taken from 120°C (250°F) cured laminate, showing slick surface with no transfer of nylon (mag X450).
Fig. 29.27 Extreme enlargement of interlaminar fracture in top layer of carbon-epoxy laminate (mag x1100).
Commentary 683 removal of a tear ply that adhered to the matrix as tenaciously as conventional peel plies refuse to adhere. This, in turn, raises questions about the feasibility of tear plies as a surface-preparation technique since the laminate is damaged far more than by light grit-blasting, as a comparison between Figs. 29.10 and 29.27 will attest. 29.5 COMMENTARY
In the mid-l980s, a major effort was aimed at the resolution of a problem associated with adhesive bonding of aluminum alloy structures. In that case, as in the problem discussed here, the glue failed to stick and interfacial failures resulted. However, there remains a very great difference between the two cases. In the earlier case, every factor associated with the weak bonds that passed all ultrasonic inspections was identified in a matter of weeks. The conditions had been replicated in a laboratory and the investigation closed within a few months. Significantly, there were no loose ends and the problem has not recurred. The primary cause of the problem was condensate on adhesive film that had been removed from storage before it had thawed out. However, there was a second factor involved as well. The first violation of proper processing procedures would not cause a defective bond unless the moisture was trapped at the interface between the details. This condition happened consistently with one kind of bonding tool, while similar parts made concurrently on a different kind of tool showed no such problems. The second kind of tool permitted complete ventilation of any trapped volatiles as well as of any generated during the cure. There were no large area defective bonds. Both tools produced local bond defects where a tool or the parts were out of contour. These problems were eliminated by correcting the tools and by better straightening of the stiffeners. This problem was resolved very quickly, whereas the present problem with peel plies has not been solved at
the production level in over 20 years. Because the metal-bond problem was resolved so quickly, very few panels were involved. There were no in-service failures because every affected panel was identified before delivery and reinforced by rivets. Peel-ply ’surface preparation’ for bonding of composites has not been as thoroughly explained. There have been many instances of such weak bonds not being detected until they had split apart in service, even though there is no reason to believe that the bonds wore out under mechanical loads. The observed modes of failure are consistent with a manufacturing problem. There is little doubt that gentle gritblasting is the most reliable method of preparing thermoset composite surfaces for bonding. Sanding can work only on fairly fineweave cloths in composite laminates. Otherwise, it is not possible to sand to the bottom of all depressions left by the peel ply without also sanding significantly into the structural fibers. A few other manufacturers have used tear plies, which leave a completely fractured resin surface, rather than the more commonly used easily removed peel plies discussed here. More use should be made of tear plies. However, if a tear ply were used on a lightweight honeycomb or foam sandwich panel, there is a good chance that the core would fail instead. There can also be no doubt that when problems have occurred as a result of bonding directly to the surface left by removal of a peel ply, or one with totally inadequate sanding, they are widespread and serious. These problems have been experienced at many places. Yet, if they happened as repeatedly as the defects caused by the combination of the two factors cited above, one would have expected such a backlash against the procedure that every unreliable peel ply would have been withdrawn from the marketplace, preventing any recurrence of the problems. That has not happened so, presumably some organizations are able to bond successfully to composite structures by simply removing a peel ply.
684 Surface prqarations for ensuring that the glue will stick in bonded composite structures Regrettably it seems as if anyone doing this process successfully has no reason to investigate why he is successful, so the differences between his techniques and those that lead to trouble have remained unidentified. It seems that there is something else involved as well possibly prebond moisture or something associated with venting during the cure or possibly the breakdown and transfer of nylon peel plies cured at too high a temperature or the transfer of a release agent on some peel plies. Weak interfacial bonds between the adhesive layer and both composite adherends should never have been strong enough to rip a properly cured adhesive back and forth from one interface to the other in the manner that is so evident in Fig. 29.8. And, more significantly, attempts to replicate the weak bonds in the laboratory have been inconclusive and, at times, inconsistent. This merely reinforces the assumption that the problem is not yet fully understood without, in any degree, diminishing the conviction that the problem is serious and needs to be resolved. It would be helpful if there were a reliable a peel-dominated quality-control test for composite surface preparation that was equivalent to the wedge-crack test used for metal bonding. Unfortunately, the experiments performed to date have been bedevilled with extraneous influences that are not yet understood but have a much greater effect on the strength of the coupon than the variations in surface treatment. Neverless the goal remains. 29.6 CONCLUSIONS
As acknowledged above, more technical information needs to be uncovered about adhesive bonding to fiber-polymer composite surfaces created by simply removing a peel ply. The weak bonds associated with this technique are a financial burden that is both serious and easy to avoid. The use of low-pressure grit-blasting as the final step in the preparation of these surfaces
for bonding has been found to be extremely reliable in service. While one needs to buy appropriate equipment, the cost of doing so is a small fraction of the typical cost of even one composite detail. With the right equipment, training of technicians is straightforward. There can be no valid argument in favor of not mastering the art of grit-blasting. While there may be choices for the surface preparation during initial fabrication, peel plies cannot possibly be stripped off a second time to create a new 'clean' surface during repairs. It may be that some manufacturers make successful bonds to peel-ply composite surfaces already. It may also be possible that coupling agents may be found to enable others to do the same. However, it is undeniably true that, for at least two decades, some manufacturers of composite components who have relied on peel plies alone as surface preparation for bonding have created weak bonds that have fallen apart in service. The cost of unanticipated repairs and investigations has greatly exceeded any expected initial cost savings. Given the widespread nature of these problems, it is appropriate to recommend that the use of peel-ply surface preparation alone be discontinued unless it can be shown that its use never results in interfacial failures between the composite and adhesive layers. It is quite clear, from experience, that the cost of even one in-service bond separation exceeds by far the savings derived during an entire production run by not thoroughly abrading the surface. Scanning electron microscope images, such as those presented here, show easily distinguishable differences between the matt rough surface created by mechanical abrasion and the microscopically smooth furrows associated with peel-ply removal alone. It is not difficult to ensure that the surface preparation is adequate for bonding or painting before the bond has been made. Conversely, it is extremely difficult to detect a weak bond nondestructively once the error has been made.
References 685 REFERENCES 1. Schliekelmann, R.J., Adhesive bonding and composites, Progress in Science and Engineering of Composites. In Vol. 1, Proc. 4th Intl Conf. Composite Materials (Tokyo), 22-28 Oct., 1982. Japan Society for Composite Materials and The Metallurgical Society (TMS) of AIME; (T. Hayashi, K. Kawata and S. Umekawa, eds) Amsterdam: North-Holland, 1982, pp. 53-78. 2. Hart-Smith, L.J., Ochsner, R.W. and Radecky, R.L., Surface preparation of fibrous composites for adhesive bonding or painting, Douglas Service Magazine, 1984, First Quarter, pp. 12-22. 3. Hart-Smith, L.J., Ochsner, R.W. and Radecky, R.L., Surface preparation of fibrous composites for adhesive bonding or painting, Canadair Service News, 1958,14(2), 2-8. 4. Parker, B.M. and Waghom, R.M., Surface pretreatment of carbon fibre-reinforced composites for adhesive bonding, Composites, 1982, 13, 280-288.
5 . Pocius, A.V. and Wenz, R.P., Mechanical surface preparation of graphitepoxy composite for adhesive bonding, SAMPE J., 1985, Sept/Oct, 50-58. 6. Wingfield, J.R.J., Treatment of composite surfaces for adhesive bonding, Int. J. Adhesion and Adhesives, 1993,13(3),151-156. 7. Hart-Smith, L.J., Joining of organic-matrix composites. In A S M Handbook, Vol. 6: Welding, Brazing, and Soldering, Ohio: ASM Intl, Dec. 1993, pp. 1026-1036. 8. Mahoney, C.L., Fundamental factors influencing the performance of structural adhesives, Internal Report, Dexter Adhesives & Structural Materials Division, The Dexter Corporation. 9. Hart-Smith, L.J., Design and development of the first Lear Fan all-composite aircraft, Douglas Paper 8184, presented to Institution of Mechanical Engineers Conference on Advanced Composites, London, England, March 7-8, 1989.
LAMINATE DESIGN
30
Jocelyn M . Seng
birch or spruce, laid over balsa core or fir stringers to form a sandwich structure1. This An early example of laminated composite successful production aircraft (7781 units built) materials is the de Havilland Mosquito was designed without the analytical techniques fighter/bomber used by the British Royal Air described in this chapter and without fancy Force during World War I1 (Fig. 30.1). This aircomputer tools. With current laminate design craft was built entirely out of wood because of and analysis techniques, today’s higher-perforlimited metal supplies and the need for quick mance composite aircraft are made possible; delivery. The wings, for example, were made as with the increase in speed and accuracy of comthree-ply skins (each 1.5 mm (0.060 in) thick) of putation results, designer confidence in 30.1 INTRODUCTION
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Fig. 30.1 The de Havilland Mosquito, an all-wood WWII Production aircraft.Printed With Permission, Zokeisha Publications, Inc.
Laminated plate theory 687 composites structures is increased. This chapter presents the basic mathematical tools used to design laminates and provides insight on the many options for optimizing the material for particular needs. Three distinct levels of benefit can be derived when using composites and laminate design. With equal fiber distribution in multiple directions, rendering an effectively quasi-isotropic material, composites can approximate metals while providing a weight savings due to the difference in material densities. In addition, however, designers opting to use composite parts enjoy the advantage of being able to tailor the properties of their material by orienting load-carrying fibers in the directions that there are loads. The result is an anisotropic material, which by definition is a material with different properties in different directions. Ultimately, the composites industry is finally beginning to see the development of unique structures that have never before been attempted and with material behavior that is only possible with distinct laminate designs. This is a result of coupled behavior, for example, an extensional load on an anisotropic material can yield extension coupled with bending and twisting deformations. The objective of this chapter is to outline a method to design ply layouts which achieve structural design goals for composite parts. This method is based on laminated plate theory used with the quadratic failure criterion2. This general discussion, which assumes familiarity with undergraduate mechanical engineering fundamentals, shows how the principles for isotropic materials (such as metals) are extended to the analysis of advanced composites. The basic equations are presented and the analysis procedure is outlined. Simplifying concepts are introduced and discussed. Simple computer codes that embody these equations are now widely available, making it unnecessary to ever have to solve these equations by hand. Examples and sample problems are included to demonstrate concepts.
30.2 LAMINATED PLATE THEORY
30.2.1 LAMINA
Advanced composite materials are typically supplied as a thin layer, called a ply or lamina, which is subsequently stacked into a thin plate, called a laminate. A unidirectional ply or lamina is a flat or curved layer of fibers oriented in one direction and held together by matrix material that serves to support the fibers. The stresses perpendicular to the planar surface are assumed to be zero. While the behavior of isotropic materials can be described with two elastic constants (typically the Young’s modulus and the Poisson’s ratio) and one strength value, a composite ply that is transversely isotropic is characterized by four elastic (stiffness) constants and five strength parameters in two-dimensional analysis. The material properties are defined along the fiber (x-direction ) and perpendicular to the fiber (y-direction ). For each unidirectional ply in its own axes, the four orthotropic elastic constants are the longitudinal tensile modulus, Ex; the transverse tensile modulus, EY;the major Poisson’s ratio, vx; and the shear modulus, Es. Only one Poisson’s ratio is necessary since Y,,= vx(€,,/Ex). The five strength parameters for each unidirectional ply are the longitudinal tensile strength, X; the longitudinal compressive strength, X’; the transverse tensile strength, Y; the transverse compressivestrength, Y’; and the shear strength, S. The five initial coupon tests to experimentally determine the nine material constants are shown in Fig. 30.2. In the stress, o, versus strain, E , plot, the material is characterized by the slope of the line, which represents material stiffness, and by the failure point, which defines the maximum stress that the material can sustain, i.e. its strength, and by the corresponding maximum strain. Use of the strength parameters is discussed in the section on failure criteria. During three of the coupon tests, the nonzero strains are monitored and the relationships in eqn (30.1) are determined. The
688 Laminate design Transverse
Longitudinal
EX
"
j
=-- EY
......
Longitudinal Compression
Transverse Compression
Fig. 30.2 Coupon tests to determine the nine material constants used to characterize an anisotropic material.
material is assumed to be linear and elastic; thus, the stiffness of a material is the same in tension as in compression. Based on the four elastic constants, infinitely different laminates can be designed using laminated plate theory.
If a unidirectional specimen was simultaneously tested under the three load cases, longitudinal tension, transverse tension, and shear, then superposition of the strains results in
Longitudinal tension test 1
(30.1)
= - 0
E
Ex
--0VX
= -0
E
3
EY
Y
ES
=
Ex
Y
1 -0, E S
Written conveniently in matrix notation,
Transverse tension test 1 = - 0
E y
E,,
y
0
0
Shear test 1
E
=-a, E,
(30.3)
Laminated plate theory 689 Defining the plane stress stiffness matrix [QJ = [SI-', another form of eqn (30.3) is
l o
0
E,
laminate, x, y, s and 1,2,6 are interchangeable. Material properties are specified with respect to the on-axis coordinates. The properties of an off-axis ply, anything other than 0 degrees, can be calculated by transforming the properties of the 0-degree ply. The angle of transformation, 8, is equal to the ply angle shown in Fig. 30.3, where 1 and 2 are the laminate axes and x and y are the rotated ply axes. 8 is positive counterclockwise from the 1-axis to 90°, and negative clockwise to -90".
Y
J
This calculation of the plane stress stiffness matrix [Q]for a single ply is the starting point of laminated plate theory, once the engineering constants have been experimentally determined.
1
30.2.2 COORDINATE TRANSFORMATIONS
Two coordinate systems are used in laminated plate theory. The local, or on-axis, coordinate frame is defined by x and y, also referred to as the ply axes. The x-axis is along the longitudinal direction of the ply (along the fiber); the y-axis is in the same plane, but in the transverse direction (perpendicular to the fiber direction). The subscripts is used with expressions for shear, and is a contraction for the subscript xy. The ply's material properties are defined in this axis system. Since not all plies are aligned in the same direction along the principal loading axis in a laminate, a second set of coordinates is necessary to analyze composite laminates. The global, or off-axis, coordinate frame is defined by 1 and 2, also referred to as the laminate axes. The 1 direction is along the principal orientation of the laminate; the 2 direction is perpendicular to it. The subscript 6 is used with expressions for shear, and is a contraction for the subscript 12. The loads on the laminate and the boundary conditions are usually defined in the global svstem. In the case of a 0-degree unidirectional , "
Fig. 30.3 Definition of ply axes (x,y) and laminate axes (1,2), where the lines indicate the fiber direction.
The laminate off-axis stiffness matrix is computed from the ply on-axis stiffness matrix by using the following ransformation relation:
Q1ll Q22
m4 n4 m2n2 m2n2 m3n mn3
2rn2n' 4m2n2 2m2n2 4m2n2 m4+n4 -4m2n2 (m' -n2)' m2n2 --2m2n2 -mn3 mn3 - m3n 2(mn3- m3n: -m3n m3n- mn3 2(m3n- mn3: n4 m4 m2n2
(30.5)
690 Laminate design where rn = cos 8, n = sin 8. A summary of some other useful transformation relations is given in Section 30.7. It is clear from these relations that when 8 = 0, then Q,, = Q,, = 0, which means shear and extension are uncoupled, i.e. shear loading only causes shear deformation, and extensional loading causes extensional deformation with Poisson’s effect but no shear deformation. Using the off-axis plane stress stiffness coefficients, the constitutive relations of eqns (30.4) and (30.3) can be generalized to a ply of any orientation:
with the [Q] matrix, the S12(and SJ, terms reflect the deformation in the direction perpendicular to the direction of loading, and are commonly referred to as the ‘component due to Poisson’s effect’ (implying the major Poisson’s ratio). The S,, and S2, (and S,, and S62)terms reflect the amount of shear deformation under extensional loading, and are commonly referred to as ’coupling terms’. Unlike the off-axis unidirectional ply shown in Fig. 30.4@), the 0-degree unidirectional ply shown in Fig. 30.4(a) does not exhibit any shear deformation under extensional loading. 30.2.3 KINEMATICS
and inversely
For example, a physical interpretation is shown in Fig. 30.4. In general, the terms on the diagonal (S,,, S,, S,) reflect the amount of deformation in the direction of loading. As
Kinematics is the study of movement and depends solely on geometry, not on material properties. Since composite laminates are often thin two-dimensional structures, plate theory is used to simplify the three-dimensional behavior. Plate theory tries to account for stretching and bending behavior relative to the midplane of the laminate. The key assumption of plate theory is that normals remain normal, straight and unstretched. In practical terms, the plies in the laminate are assumed to be completely bonded to each other, allowing no interlaminar shear. Some other assumptions are that the
Fig. 30.4 Extensional loading of (a) a unidirectionalply and (b) an off-axis ply and their associated off-axis stiffness matrix.
Laminated plate theory 691 material exhibits perfectly linear elastic behavior and that there is a perfect bond between the fiber and matrix. The out-of-plane displacement, w, can be described by a function of the in-plane coordinates such that
w = wo(x,y)
E: -I-ZK1
(30.7)
Based on the Kirchhoff assumptions, Fig. 30.5 shows the deformation of a cross section of the plate in the x-z plane, relative to the x-direction
u
= uo-z-
aw0
ax
(30.8)
30.2.4 STRESS RESULTANTS
Similarly, the displacement along the y-axis is Y = Y,-Z-
awo
aY
Based on the definition of strains,
au =-=--z-
& 1
ax
auo azw, ax a x 2
Just as beam theory defines net tensile force, shear force, and moment, relating everything (30.9) to the neutral axis, plate theory defines stress resultants and moment resultants to eliminate any z-direction dependence and to relate everything to the midplane, as shown in Fig. (30.10) 30.6. Ply stresses along each loading direction are summed for the laminate: (30.12)
M-
2
Undeformed Cross Section
Deformed Cross Section
Fig. 30.5 Extension and bending deformation.
Fig. 30.6 Force and moment resultants acting on a plate.
692 Laminate design
A,, indicates the relationship between longitudinal in-plane load, N,, and the longitudinal extension, E;; A,, indicates the coupling between longitudinal in-plane load, N,, and the extension in the transverse direction, E;, (the traditional Poisson's effect); A,, indicates the coupling between longitudi+ [B](K) (30.13) nal in-plane load, N,, and the in-plane shear, E;; A, indicates the relationship between inplane shear load, N,, and the in-plane Bll 1' 2 B16 shear distortion, E;; B,, indicates the coupling between transverse 2' 1 2 '2 B26 in-plane load, N,, and the twist, K ~ ; '61 '62 '66 B, indicates the relationship between inDll D12 D16 plane shear load, N6, and the twist, K ~ ; D21 D26 D12 indicates the coupling between longitudinal bending load, MI, and the transverse D61 D62 D645 bending curvature, K,; (30.14) D, indicates the relationship between twisting moment load, M6, and the twist angle, K ~ .
These resultants can be rewritten in terms of strain by substituting in the constitutive relations. Putting the two-part expression for strains, eqn (30.11), into the constitutive relations, eqn (30.6), and substituting the resulting stress expression into the definitions for the resultants, eqn (30.12),
{N] = [A] (E'}
All
1 '1
1' 2
1 '6
B21
'22
B26
6' 1
B62
'66
The ply stacking sequence has no effect on the (30.15) A matrix coefficients, which reflect in-plane behavior. However, since the B and D matrix coefficients are a function of z, they are dependent on the stacking sequence. There are two unique physical situations that deserve mention. When the laminate is symmetric about its midplane, the B coefficients are zero, which means that there is no It can be shown that A,, = A,,, B,, = BZ1, D,,= coupling between in-plane loads and curvaDZ1,etc. Equation (30.14)represents the funda- tures, nor between bending loads and mental relationships in laminated plate theory. in-plane deformations. Another common sitThe 6 x 6 matrix is the laminate stiffness uation is when the A,, and A,, coefficients are matrix. zero (usually in the presence also of all the €3 A composite with unidirectional plies lami- coefficients being zero): this arises when a nated in different directions (a generally laminate is balanced, i.e. there are an equal anisotropic material) under an inplane load number of off-axis plies in the +8 and 4 may stretch, bend and twist, as a result of directions and they have equal thickness. In extensional/ shear coupling. By comparison, a this case, there is no coupling between extenmetal structure will stretch only under an sion loads and shear strain. If, in addition, inplane load, bend only under flexure, and those +8 and 4 plies are effectively the same twist only under torque. Each matrix coeffi- distance from the midplane, then the correcient in eqn (30.14) relates a particular sponding B and D matrix coefficients tend resultant to a strain expression. For example, toward zero.
[AI = IIQldz
Laminated plate theory 693 Most laminates used today are symmetric to eliminate or reduce any tendency of the structure to warp unexpectedly. Most laminates are also balanced, often because it is erroneously thought to be necessary to prevent the structure from warping. A balanced laminate is really only necessary in situations with reversible shear loading conditions. 30.2.5 RESULTING STRAIN STATE
Knowing the laminate stiffness matrix and the applied loads, the resulting strains can be computed. The strains are obtained by inverting the stiffness matrix and multiplying by the input load. Instead of inverting the 6 x 6 stiffness matrix, however, it is sometimes possible to simplify the analysis even further. If the laminate is symmetric about the midplane so that the B coefficients are identically zero, then the in-plane (described by N, [A], E ) and bending problems (described by M I [D], IC)become uncoupled. In this case, it is much easier to invert two 3 x 3 stiffness matrices to get the compliance matrices (see also Section 30.7 for the explicit terms to invert a matrix)
a flexural contribution, then it must be added
The off-axisply strains can be transformed to on-axis ply strains for each ply and their significance can be evaluated per a failure criterion (refer to the relations given in Section 30.7).
m2 n2 =.
[Et Es
n2 m2
-mn mn 2mn -2mn m2-n2
i I
f:
.,
I
(30.19)
where, as before, m = cos 8, n = sin 8. One step further, the on-axis ply stresses can be obtained by multiplying the on-axis ply strains by the ply stiffness matrix [Q] as shown in eqn (30.6). The laminate engineering constants, which have meaning with symmetric laminates only, are calculated from the compliance matrix and are useful for comparison to the properties of other materials, such as metals
[a] = [AI-’ [d] = [Dl-’
(30.16)
Then, the compliance matrix is multiplied by the appropriate input load conditions to compute the laminate strains ‘EP E; \E:
=
I
‘11
‘12
‘16
‘IN,
’
a12 aZ a26 N2
‘16
‘26
‘66. tN6
(30.17)
where the compliance terms have been normalized to have the necessary units of [length2/ force] [a*] = [a]h
(30.21)
,
In summary, the mathematical process of analyzing composite laminates is indicated in Fig. 30.7. A laminate’s stiffness is calculated as the summation of its individual ply properties. Load on the laminate is described in terms of Since all the plies are bonded together, the the laminate coordinates. Calculated from the strains in each ply, in the laminate axes, (1,2), applied load and the known material stiffness are equivalent to the laminate strain. If there is properties, the response of the laminate is
694 Laminate design
expressed as laminate strain. In order to apply a failure criterion, the laminate strain is commonly transformed into ply strains and each ply is individually evaluated. 30.3 ENHANCEMENTS TO THE BASIC
LAMINATED PLATE EQUATIONS
30.3.1 SANDWICH CONSTRUCTION
In composite structures, sandwich constructions are commonly used. By increasing the distance between the load-carrying laminate skins, a core can provide increased bending stiffness without a significant weight penalty. The core is often idealized in laminate design: it is assumed that the core does not contribute to laminate strength or in-plane stiffness, and that the shear bonds between the skins and core are perfect. The parallel axis theorem can be used to account for the increased moment of inertia that the core creates by offsetting the laminate load-carrying skins from the midplane. Once the skin laminates have been sized, further calculations can be performed to confirm that the core assumptions are valid3.
Measured
Ply Stiffnesses
Plane Stress Coeff.
Laminate Stiffness Matrix
30.3.2 HYGROTHERMAL EFFECTS
Most structure is exposed to a variety of environmental effects. Of particular concern are heat and moisture. The design must account for the hygrothermal effects (hygrothermal means water and temperature). A laminate that is stress-free when curing at an elevated temperature will have residual stresses when brought back to room temperature. It has been thought that some of the apparent improvement in toughness of 250°F resins over 350°F resins is simply due to the reduction in residual stresses (AT = 75 - 250 = -175°F versus AT = 75 - 350 = -275°F). In addition, moisture is absorbed by the laminate, usually into the resin. The negative effect on the mechanical properties is particularly pronounced at both high temperature and high humidity. Assuming that the primary effect on residual stresses is due to different thermal expansions and moisture expansions along and transverse to the fiber direction, two additional strains on the laminate result:
Inverted Stiffness Matrix
{E)
= {a)AT+@)Ac
Given Loads; Resulting Laminate Strains
(30.22)
Ply Strains
Fig. 30.7 Logical flow of calculations involved in analyzing a symmetric composite laminate that is loaded axially and/or in bending.
Failure criteria 695 The summation process can be used to determine the effective laminate expansion coefficients. The hygrothermal load, sometimes called non-mechanical load, can be computed by multiplying the laminate stiffness by the hygrothermal strain (i.e. laminate thermal expansion coefficients multiplied by the change in temperature). The stress induced by moisture absorption can be accounted for similarly by using PAC in place of aAT. Thus, the non-mechanical loading in the laminate can be expressed as
The mechanical and non-mechanical loads, N and W , can be added together to determine the total load experienced by the laminate. 30.4 FAILURE CRITERIA
The ultimate objective in any structural design is to create a structure able to withstand deflections or loads without failing. The initial concern is to remain below a prescribed deflection as part of stiffness criteria. Once these criteria are satisfied, the focus shifts to a strength criterion, such that applied stress must not exceed laminate strength. Composite materials normally possess different strengths when loaded in either tension or compression. The following represent the minimum number of strength properties necessary to characterize a unidirectional or fabric ply. They are determined using material coupon tests, previously outlined in Fig. 30.2.
X = longitudinal tensile strength,
X' = transverse tensile strength, Y = longitudinal compressive strength, Y' = transverse compressive strength, S = shear strength. (30.24)
The maximum ply strain values can be interpreted by dividing the above strengths by the appropriate ply stiffness coefficient. cx* = E ~ ' *= E * = Eyf* = Y E,* =
max longitudinal tensile strain, max longitudinal compressive strength, max transverse tensile strength, max transverse compressive strength, max shear strength. (30.25)
Laminate strength is function of material (ply) strength and the constraints on the ply within the laminate. Thus, failure is best assessed at the ply level. The proper interpretation, however, of the significance of the applied stress relative to the material strength is still debated. Maximum stress and maximum strain failure criteria are common wherein the applied stress or strain value is compared directly to the strength value. A review of failure criteria has been published4. Early laminate failure theories fail to account for Poisson's effects and interaction between loads in orthogonal directions (a complex load condition). For example, the major weakness of both the maximum stress and the maximum strain failure criteria are their inability to couple stress, or strain, components in determining the ultimate failure of a ply. It is important to understand that the longitudinal tensile failure of a ply is affected not only by the longitudinal load, but also by the magnitude of applied transverse loads. As a result, stress interaction criteria are widely used throughout the industry to determine ply failure in a laminate.
30.4.1 QUADRATIC FAILURE CRITERION
Tsai developed a two-dimensional stress interaction failure criterion and predicted the strength of an orthotropic ply subjected to combined stresses or strains. This analysis
696 Laminate design
takes into account the effects of other stress components on the strength in any one direction. Tsai postulated a criterion in stress space consisting of the sum of linear and quadratic scalar products as follows:
+ Ftp1I 1 i, j
F ‘1DOI 1
=
x, y, s
(30.26)
or, in expanded form,
Equation (30.28) is substituted into eqn (30.26) and the solution of this quadratic equation can be obtained.
+ ( F p x + Fyuy)I1 Note that Fxs = Fys= F, = 0, and the six strength parameters are interpreted from the ply strength values (reviewed in Section 30.8):
F = -
1 s 2
F =
1
1
-- -
x x
before failure occurs); R < l failure has already occurred (i.e. occurs prematurely at some point below the applied stress or strain) and the applied stress or strain level can not be attained (e.g. if R = 0.5, then only half the applied stress can be sustained).
F = y
1
1
-- -
Y
Y’
(30.27)
The stress interaction term, Fx,*, can have a value of -1 I Fx; < 1although is recommended to be -1/2. When F,; = -1/2, the quadratic failure criterion is a general case of the von Mises criterion (Section 30.8.2). Instead of simply evaluating the failure criterion to determine if the laminate failed, it is useful to consider a nondimensional ratio to provide a perspective of the significance of the applied stress relative to material strength. Tsai defines the strength ratio, R, such that
(30.28)
The strength ratio is always a positive number with the following physical implications:
X = 1 failure occurs (at the applied stress or strain level); R > 1 failure has not occurred and R represents a factor of safety (e.g. if R = 2, then the applied stress can be safely doubled
[ F I , ~ l ~ ~+][F,u,] R‘ R -1 = 0
i, j
(30.29)
= x, y, s
The positive and negative roots of the quadratic equation can be found and represent failure of the laminate in tension and compression (where the absolute value of the negative root is used), respectively. Failure envelopes can be plotted to show laminate strength for any combination of loads. Instead of the stress space representation, however, the examination of failure envelopes in strain space is a useful alternative. The representation of failure envelopes in strain space is preferred because strain is usually specified in laminated plate theory. Strain, unlike stress, is at most a linear function of the thickness. Furthermore, failure envelopes are fixed in strain space, and are independent of other plies with different angles which may exist in a laminate. Thus, they can be regarded as material properties. Another additional advantage of strain space is that the axes are dimensionless. 30.4.2 STRENGTH OF LAMINATES
Traditional failure criteria based on strength of materials are limited to the prediction of the FPF, the point beyond which the continuous and homogeneous material assumptions are no longer valid. The use of a simple method for modeling of degraded plies is recommended, from which the FPF can be estimated. The load-carrying capability of a laminate
Laminate design 697 beyond the FPF can be formulated using a ply degradation model. Two possible methods are recommended: first, the simplified micromechanics model based on the modified rule-of-mixtures relations can be used. Plies with transverse cracks are replaced by plies with reduced matrix modulus, Em. Micromechanics translates the effect of the altered constituent material properties to the ply level, e.g. how a change in the matrix modulus affects the shear and transverse modulus of the unidirectional ply. Degraded plies are modeled by quasi-homogeneous plies so that laminated plate theory can be reapplied to determine the ply stresses and ply strains. Another approach for the prediction of post-FPF strength can be based on macromechanics, without resorting to micromechanics. The degradation factor (DF) is applied directly to the transverse and shear modulus, as well as the major Poisson's ratio. The exact value for the degradation factor must be determined empirically.A value between 0.1 and 0.3 is recommended. If the degradation factor is given a value close to zero, the quadratic failure criterion can be made to resemble the maximum strain criterion and results in a generally conservative estimation of laminate strength.
of the preselected orientations results in a quasi-isotropic laminate. This is the performance baseline, because load-carrying fiber is in effectively all directions. Laminate performance can only be improved beyond that of a quasi-isotropic laminate as fiber is biased into load directions, since, of course, fiber would never be put in unnecessary directions. Heretofore, quasi-isotropic laminates have been used because they give properties like those of metals, and predictable responses that are familiar, although they are not optimal in strength-to-weight or stiffness-to-weight ratios. Many laminates used today on aircraft structures tend to be of this type. In general, however, the more directional the loading, the bigger the payoff possible with anisotropic tailoring. To improve on the performance obtained with a quasi-isotropic laminate, the cost to design and analyze the anisotropic part (using the tools like those discussed in this chapter) is unfortunately often thought not to be worth the additional weight savings. This attitude is commonly rationalized by worry about holes, increase in work associated with more complicated fiber placement (preform assembly), etc. In practice, laminate designs, if not quasiisotropic, are certainly still symmetric about the midplane, balanced (equal quantity of -8 30.5 LAMINATE DESIGN and +8 plies), and orthotropic. Capitalizing on To simplify the analysis, it is commonly initially the benefits of anisotropy will probably occur specified that a laminate will be constructed of in other industries first before being adopted plies oriented with fibers in a few preselected by the more conservative aircraft industry. directions, where only the percentage distribuAn exception to traditional aircraft laminate tion in each orientation must then be design is the X-29 experimental aircraft, which determined. Laminates with plies distributed demonstrated a unique attribute of anisotropy every 45" are called n/4 laminates (plies can be (Fig. 30.8). The basis for this design lies in the in the 0, 45, 90 or 4 5 directions. Ply orienta- important assumption that the 1,2,6 axes are tions are usually specified as a value between usually the primary load directions for the -90 and 90". For example, instead of identifying laminate. With the coordinate system for loadthe orientation as 135, the laminate orientation ing changed to be 20" off a designated is more commonly called 45", although they laminate system, it can be shown that the lamare the same). Another class of laminates are inate behavior in flexure and torsion is called n/3, where plies are placed every 60" coupled. In fact, twisting will result with flex(plies can be in the 0, 60 or -60 directions). In ural loading, even though the material would both cases, an equal percentage of plies in each normally behave as most metals. This is the
698 Laminate design
principle used on the X-296. The normal ten- the laminate. Composite materials are not dency for forward swept wings to diverge at merely a light-weight substitute for heavyincreasing speeds was counteracted by this weight metals. Structural performances which laminate design: the increase in lift creates a are not possible with metals are easily achievdecrease in angle of attack, as the laminate able. Examples of such unique properties twists in the direction opposing the forces. include Poisson’s ratios greater than unity or It is conceivable that in the future the even negative, bending-twisting coupling, and graphite golf shafts currently gaining in popu- zero or negative coefficients of thermal expanlarity could be tailored to the individual golfer. sion (CTE). The problems and examples below The same coupling principle could be applied. illustrate the engineering constants of angleA golfer’s tendency to consistently slice the ply and related laminates. Examples of large ball might allow the designer to customize a and negative Poisson’s ratios and examples of golf shaft which not only bends, but also bend-twist coupling are also given. twists slightly under the bending load of the bad swing. 30.5.2 UNUSUAL POISSON’S RATIOS
30.5.1 UMQUE BEHAVIOR
The most unique features of composite materials are the highly direction-dependent properties. Highly coupled deformation and load-carrying capability can be designed into
Personal computer software based on a computer spreadsheet allows rapid sensitivity studies and parametric analysis of the behavior of laminates. Laminated plate theory with micromechanics is programmed into ’MicMac/In-Plane’2. A companion charting tool,
Fig. 30.8 Top view of the Grumman X-29 aircraft with wings that twist under flexure to counteract the detrimental aerodynamic effects.@ NASA)
Laminate design 699 'Chart-quick', can be used to plot variation of CTE as a function of independent variables (0, E,,, E,, vf, etc.). For the following problems and examples, the carbon fiber reinforced polymer material data used are shown in Table 30.1. Figure 30.9(a) shows the engineering constants for a unidirectional laminate as it is rotated from the on-axis. The Poisson's ratio, vx, of a 0" laminate is approximately 0.3. With increasing angle of the off-axis laminate, the Poisson's ratio decreases. The Poisson's ratio of a 90" laminate is effectively zero, because contraction in the transverse direction is constrained by the fibers. Figure 30.9(b) shows the engineering constants for an angle-ply laminate. It is interesting to observe the very large Poisson's ratio of 1.32 for a [ S O ] laminate. A value of greater than one implies that the transverse dimensional change is more than in the dimensional change in the longitudinal direction of loading. When the ply angle is either 0 or 90", the laminates (and consequently the values for the engineering constants) in Figs. 30.9(a) and 30.9(b) are the same. In both Figs. 30.9(a) and 30.9(b), the transverse modulus, E,, is a 'mirror
Table 30.1 Material property data for three different carbon fiber systems: IM6/Epoxy, T300/5208 and M40J/F584 lM6/ Epoxy
T300/
5208
M40J/ F854
Longitudinal tensile modulus, E x (Msi)
29.44
26.27
32.8
Transverse modulus, EY(Msi)
1.62
1.49
1.2
Poisson's ratio
0.32
0.28
0.26
Shear modulus, Es (Msi)
1.22
1.04
0.66
Longitudinal CTE, a1 Transverse CTE, a2
15
Volume fraction V,(%)
66
70
22.50
1.5 22.50
15.00
1 0.5
7.50
0.00
0
0.00
f
Poisson's Ratio 2 1.5
15.00
7.50
Ply Angle, 8 (degrees)
62
image' of the longitudinal modulus, Ex. Figure 30.10(a) shows the engineering constants for cross-ply laminates. For any given laminate, the longitudinal modulus, Ex, and the transverse modulus, E , are equal. The Poisson's ratio, vx, of a [d/90] laminate is approximately zero, because of the presence of fibers in the transverse direction. The largest
Modulus Poisson's (Msi) Ratio 30.00 2
Modulus (Msi) 30 00
(4
-0.14
-0.5
(b)
1
'
0.5 0 15
30
45
60
75
Ply Angle, 8 (degrees)
Fig. 30.9 Engineering constants of IM6/epoxy laminates as a function of 6 for (a) off-axis unidirectional and (b) mgle-ply [+el,.
[e],;
700 Laminate design
LO7 (6 + 90>1,,
LO,
f
6,Is
Modulus
Poisson's
(Msi) 30.00 22.50 15.00 7.50
0.00 15
30
45
60
15
75
Ply Angle, 8 (degrees)
(4
30
45
60
75
Ply Angle, 8 (degrees)
(b)
Fig. 30.10 Engineering constants of IMG/epoxy laminates as a function of 19for (a) cross-ply [I9,(0+ 90)],,;
and @) LO,,
* qs.
Poisson's ratio is 0.55 for a [*45] laminate. The shear modulus, E , is a maximum, of course, for the [*45] laminate. Figure 30.10(b) shows the engineering constants for laminates with 50% 0" plies and 50% angle-plies. With the exception of the transverse modulus, the results are similar to those for the angle-ply laminate shown in Fig. 30.9(b).When the ply angle is 90", the values for the engineering constants in Figs. 30.10(a) and 30.10(b)are the same. Figures 30.11(a) and 30.11(b)show the engineering constants for some unusual laminates. When the ply angle, 8, is 15", Fig. 30.11(a) shows an off-axisunidirectional laminate and Fig. 30.11(b)shows an angle-ply. For all other ply angles, the laminates are unbalanced. From Fig. 30.11(a),it can be observed that the [15/60Is laminate exhibits an extremely large negative Poisson's ratio of -0.32, meaning the laminate will expand in the transverse direction under longitudinal tension loading and compress in the transverse direction under longitudinal compressive loading. From Fig. 30.11(b), it can be observed that the [-15/30Is
laminate exhibits a very large Poisson's ratio of 1.32, when compared with that of an isotropic material (0.3). Besides the unique Poisson's ratio behavior, it is also important to examine the values of the other coupling coefficients.
EXAMPLE Table 30.2 considers the resulting deformations on coupon specimens under load, and Fig. 30.12 indicates the relative magnitude of deformation due to large and negative Poisson's ratios. 30.5.3 STIFFNESS AND COUPLING
It is useful to look at the A, B, D stiffness matrices of some simple laminates. For ease of comparison, the stiffness matrices can be normalized to have units of [force/length2]by defining
[A*]= [ A ] / h , [B*] = 2[B]/h2,
Laminate design 701 Poisson‘s Ratio T 2
,r”
30.00 T ’ 22.50
[-I 5/9 14s
Modulus (Msi) 30’00
Poisson‘s Ratio
T
22.50
1.5
15.00
1
1
15.00 7.50
0.5
7.50
0.5
0.00
0
0.00
0
1
Ply Angle, 6 (degrees)
1 -0.5
(4
Ply Angle, 8 (degrees)
-0.5
(b)
Fig. 30.11 Engineering constants of IM6/epoxy laminates as a function of 8 for (a) [15/8],s; and (b) [-w~I,.
Table 30.2 Strains, deformations and strength ratio (based on first-ply-failure) of 10 in x 1 in x 0.1 in specimens under 1000 lb longitudinal load, N, Longitudinal strain Material
Transverse strain E2
El
(1C3in/in)
in/in)
Longitudinal Transverse displacement displacement A1
A2
(1C3in)
(10-3 in)
3.4 3.4 6.5 9.5 10.2 10.9 24.0 32.0 62.5
-1.0 -1.1 -0.2 -12.9 -3.0 3.5 -17.7 -20.7 -1.1
Strength ratio R
.-
40ksi Steel IM6/Epoxy IM6/Ep IM6/Ep 30ksi Aluminum IM6/Ep IM6/Ep E-glass/Ep IM6/Ep
0.34 0.34 0.65 0.95 1.02 1.09 2.40 3.20 6.25
-0.10 -0.11 -0.02 -1.29 -0.30 0.35 -1.77 -2.07 -0.11
3.9 50.0 7.9 6.8 2.9 3.1 2.7 4.5 0.8
Fig. 30.12 Relative deformation of 10 x 1 x 0.1 in specimens under 1000 lb load along the centerline (laminates are IM6/epoxy, unless otherwise indicated).
702 Laminate design
B* matrix with nonzero terms. The first and fourth laminates are balanced and so the A*16 and A*26coefficients are zero. For the second and third laminates which differ by the sign of the off-axis plies, the stiffness behavior differs only in that the A*16,A*26,D*,6 and D*26coefficients are of opposite signs. Table 30.5 displays different quasi-isotropic laminates. Note that the normalized A* matrix
A four-ply laminate consisting of two 0" and two 90" plies can be combined into four different laminates. From Table 30.3 it can be observed that while the A* matrix remains unchanged through varied ply stacking sequences, large differences arise in the B* and D*matrices. From Table 30.4 it can be observed that only the fourth laminate is unsymmetric and has a
Table 30.3 Normalized stiffness coefficients for four IM6/epoxy laminates, in units of Msi 190/0/0/901
10/90/90/01
Layup
' 15.624
[A*]
.15.624 0.525
0 ' 0.525 15.634 0 0 0 1.220,
0.525
0.525 35.634 0 0 0 1.220
~
0
0
0
0
0 ' 0
, o
0
o , ,
'
[B*]
26.129 0.525 0.525 5.138 0 0
0 0
.
1.220.
'
0
0
0
0
0
0
' 5.138
.
0.525 0.525 26.129
0
0
f 0/0/90/901
~0/90/0/901 '
15.624 0.525
15.624 0.525 0.525 15.634
0
,
:]I0
0
3.499 0 0 -3.499
0
0 ' 0 1.220,
0
' 15.634 0.525 0.525 15.634
0
,
0.525 15.634
0 1.220
0
0
0 -
- 6.997 0 0
0
o * , 0 0
0
0
1.220
0
0 ' 0 0 .
-6.997
0
. .15.634 0.525
1.220,
0 ' 0.525 15.634 0 0 0 1.220,
Table 30.4 Normalized stiffness coefficients for four IM6/epoxy laminates, in units of Msi Layup
[0/0/+45/-451$
10/0/+45/+451+
' 19.463 3.692 [A*]
,
0 ' 3.692 5.469 0 0 4.387. o
* o IB"1
0
0 0
, o
0
:I 0
27.087 1.316 0.656 1.316 2.597 0.656 0.656 0.656 2.012
' 19.463 3.692 3.499 3.692 5.469 3.499 , 3.499 3.499 4.387
0 0
J
o
0
0 0
[0/0/-45/-45Is
I
19.463 3.692 -3.4993.692 5.469 -3.499 .-3.499 -3.499 4.387
:ll! 0
27.087 1.316 0.875 1.316 2.597 0.875 0.875 0.875 2.012
f+45,/45,1
0 0 0
.
0 6.859 9.297 0 , 0 0 7.555 * 9.297 6.859
: [:
0 3.499 0 3.499 3.499 3.499 0
0
27.087 1.316 -0.875 1.316 2.597 -0.875 -0.875 -0.875 2.012
II
9.297 6.859 6.859 9.297 0
0
0 0
7.555
I
Laminate design 703 Table 30.5 Normalized stiffness coefficients for four IM6/epoxy laminates, in units of Msi
0 12.466 3.692 3.692 12.466 0 0 0 4.387
20.932 3.098 1.312 3.098 5.188 1.312 1.312 1.312 3.793
I I
12.466 3.692 0 3.692 12.466 0 0 4.387 0
1[
20.2115 2.504 1.968 2.504 7.094 1.968 1.968 1.968 3.200
12.466 3.692 0 3.692 12.466 0 0 0 4.387
12.466 3.692 0 3.692 12.466 0 I O 0 4.387
18.194 2.782 3.735 2.782 8.558 2.208 3.735 2.208 3.477
22.001 1.932 0.737 1.932 6.451 1.956 0.737 1.956 2.628
terms are equivalent for all quasi-isotropic laminates. This means all have the same stiffness to weight ratio. The differences between these laminates thus manifests themselves only in how they behave in bending.
(iii) Evaluate the displacements by integrating the strains. (Note that for the tube,
PROBLEM
1
Find the amount that an anisotropic 20-layer [0/30], T300/5208, 3 in diameter, 12 in tube tube will extend, change in circumference, and twist under an in-plane load, N, = 100 lb/in.
Multiplying out the strains, a12N2/
= a16N1
‘6
E
I
= a , , ~ , , E2 =
dd) dx
=r-.)
Eldx = l:$dx
= l:llNldx
E2dy=$$dC
=[a12NldC
-
u = allNIL
---t
z,
= ul2N12nR
SOLUTION (i) Compute the laminate stiffness matrix and invert to get the compliance matrix:
I
21.159 2.567 3.935 2.567 2.476 1.458 , 3.935 1.458 3.195
I
62.563 -26.647 -64.893 [a*] = -26.647 563.836 -224.518 -64.893 -224.518 495.374
L
1 R
-
+@ = a , 6 ~ l R
(iv) Evaluate the displacements numerically: =-a*ll N L = - 62.563 x 100 x 12 h 0.1
I
= 0.75 x 10-3in
v = -Nl2nR a*,, h
= -___ 26*M7 x 100 x 2 x 3.14 x 1.5
0.1 = -0.25 x 10-3 in
(ii) Evaluate the strains: @
= -U*16 Nh
L = ‘R
= -0.52 x
64 893L x 100 x 12/1.5
0.1 radian = -0.03 degree
704 Laminate design 30.5.4 CTE BEHAVIOR
the a1 is -2 to -2.5 (versus -0.14 for a1 of the The following four figures show the coefficient unidirectional tape, as shown in Fig. 30.13(a)). As the number of ply angles increases, the of thermal expansion (CTE) in two principal directions (referred to as a , and a,) for CTE behavior becomes less intuitive. Figure M40J/F584 carbon fiber laminates. Figure 30.14(a) shows the CTE of a laminate with 50% 30.13(a)shows the CTE of an off-axis unidirec- 0 plies and 50% angle-ply; Fig. 30.1303) shows tional ply; Fig. 30.1303) shows the CTE of an the CTE of a laminate with 25% 0 plies, 25% 90 angle-ply laminate. From Fig. 30.13@), it can plies and 50% angle-ply. Note that when & be observed that, due to the Poisson coupling equals 45", the resulting laminate is quasieffect, laminate CTE values less than that of a isotropic [0, 90, &45] as confirmed by a , unidirectional material are possible for specific equaling a*. Examination of the fundamental ply angles. In Fig. 30.13(b), for 0 of GO to do", trends in Figs. 30.13 and 30.14 indicates
16.00
16.00 14.00
14.00
12.00
1200
10.00
10.00
8.00
am
-
6.00
600
4.00
4.00
c
200
200
0 C
a, ._ 0
0
0.00 -200
-r
0.00
p
10
20
30
40
SO
60
70
80
90 -200
0
Ply Angle, 0 (degrees)
-4.00
(b)
Fig. 30.13 Coefficient of thermal expansion of M40J/F584 laminates as a function of 8 for (a) off-axis undirectional [e,],; and (b) angle-ply [&J,.
4.00 0 c
c
200
--Ply Angle, B (degrees)
Ply Angle, 0 (degrees)
(4
(b)
Fig. 30.14 Coefficient of thermal expansion M40J/F584 laminates as a function of 0 in the following laminates (a>[o,, and (b) IO, ,90,, 4 1 , .
Laminate design 705 potential near-zero CTE laminates, particularly useful in spacecraft applications to minimize deformation due to the large cyclic thermal loading.
EXAMPLE To remove a composite shaft from a metal mandrel after elevated temperature cure, the laminate CTE in the hoop direction of the cylindrical section has to be less than that of the mandrel material to prevent lock-on. The composite is considered to be stress-free at cure-temperature, and thus the temperature loading is associated with the temperature decrease to room-temperature.
dominantly in the longitudinal direction to accommodate flexural loading (like a mast or golf club), Fig. 30.14(a) indicates that a [O,, G0JS with a steel mandrel would be a problematic choice, resulting in a composite shaft locked on to the mandrel as shown in Fig. 30.15@).There is a need for sufficient fibers in the hoop direction (90") to result in a laminate CTE less than that of the mandrel material. The CTE of the metal materials given in Table 30.6 indicates that it is easier to remove a composite shaft from an aluminum mandrel than from a steel mandrel.
Table 30.6 Average coefficients of linear thermal expansion of selected materials
Figure 30.15(a) illustrates that it is preferable Aluminum alloy to have the metal mandrel contract more than Concrete the composite during cool-down, which Invar Steel means that the metal CTE must be more than Titanium alloy the composite CTE. For a shaft with fibers pre-
a composite ' aofmandrel (significantamount fibers in hoop direction)
12.8 6.7 0.39 6.5 4.9
a composite ' a mandrel (predominantly longitudinalfibers)
comDosite
AT
Fig. 30.15 Result of different coefficients of thermal expansion (CTE), a,for metal and composite cylindrical sections under two thermal load cases.
706 Laminate design
EXAMPLE
freedom to design the laminated material is at the price of more complicated analysis. This chapter presented an explanation of the basic mathematical tools used to design laminates including laminated plate theory and the quadratic failure criterion. Some examples were provided to lend insight on the many options for optimizing the material to particular needs.
Carbon fiber room temperature curing materials are often used for composite bicycle frames. To prevent a metal-composite joint from coming apart under in-service thermal loading, the metal lug is generally on the inside. Ideally, under an in-service temperature increase due to sun shining on a part (which can be up to 82"C(18OoF)),or friction, the lug will expand more than the composites. REFERENCES This 'seizing' action will prevent the lug from 1. Harper, R. The Fifth Halford Memorial Lecture, breaking loose. If there are mostly longitudinal J. Roy. Aeronaut. Soc., Apr 1966,70,477-486. fibers due to flexural design considerations, 2. Tsai, S.W. 1990 Composites Design, 4th Edn., the CTE in the hoop direction may be greater Think Composites, Dayton, OH, 1988. than that of the metal; the metal fitting should 3. MIL-HDBK-23A, Military Handbook be then on the outside. More commonly, howStructural Sandwich Composites, Department ever, the composite has more fiber wound of Defense, Washington., 1968. 4. Quinn B.J. and Sun, C.T. A critical evaluation of around the lug (in the hoop direction) and the failure analysis methods for composite lamiCTE is less than that of the metal. These are only two of the challenges in coping with different thermal expansion coefficients of dissimilar materials. In general, because of the low CTE of the fiber, multidirectional laminates typically have a laminate CTE less than metals. This can cause difficulties in part manufacture, in cases of elevated cure. It is generally desirable to match the CTE of a composite part with the tooling to minimize springback and residual stresses. A tool with zero CTE results in a dimensionally more accurate part; a tool with equivalent CTE results in no chance of lock-on. Composite tooling systems may be used in order to match CTE; however, such tools usually do not have the necessary durability over long production runs.Steel tooling is often preferred over aluminum tooling because of the lesser CTE. 30.6 CONCLUSIONS
Composite materials have many well-known advantages over other structural materials including increased strength-to-weight ratio, stiffness-to-weight ratio, increased fatigue life, corrosion resistance, and the ability to tailor the properties of the material. The extraordinary
nates, Proceedings of the Ninth DoD/NASA/FAA Conference on Fibrous Composites in Structural Design, Lake Tahoe, Nevada, Nov 67,1991, pp. V21-V37. 5. Tsai, S.W. and Hahn, H.T. Introduction to Composite Materials. Technomic Publishing Co., 1980. 6. Hadcock, R.N. X-29 Composite Wing, AIAA Evaluation of Aircraft/Aerospace Structure and Materials Symposium, Air Force Museum, Wright-Patt AFB, Apr 24-25,1985. 7. Popov, E.P. Mechanics of Materials, 2nd Edn., Prentice-Hall, Inc., 1976, p. 295.
OTHER READING
MIL-HDBK-17-2C, Military Handbook - Polymer Matrix Composites, Vol. II: Material Properties, Department of Defense, Washington D.C., Feb 25, 1994. MIL-HDBK-17-3D, Military Handbook - Polymer Matrix Composites, Vol. 111: Utilization of Data, Department of Defense, Washington DC, Feb 28,1992. Hyer, M.W. Calculations of the Room-Temperature Shapes of Unsymmetric Laminates, J. Composite Materials, 1981,15,296-310. Hyer, M.W. Some observations on the cured shape of thin unsymmetric laminates, J. Composite Materials, 1981,15, 175-194.
Appendix B 707 Garfinkle, M. Twisting Smartly in the Wind, Aerospace America, American Institute of Aeronautics and Astronautics, Reston, VA, July 1994, 18-20. Smith, E.C. Vibration and Flutter of Stiff-Inplane Elastically Tailored Composite Rotor Blades, 34th AIAA/ ASME/ ASCE / AHS/ ASC Structures, Structural Dynamics and Materials Conferece, La Jolla, CA, AIAA-93-1302-CP, Apr 19-21, 1993, pgs. 26-37.
n2 m2
-mn n2 mn 2mn -2mn m2-n2 m2
=
(30.A8)
-2mn 2mn m2- n2 NOTE MIL-HDBK-23 is obsolete and no longer available. The data is being included in MIL-HDBK-17, which can be obtained from DODSSP, BLDG4D, 700 Robbins Ave., Philadelphia, PA 19111-5094, or http:/ /www.dodssp.daps.mil.
~
.
30.7 APPENDIX A: TRANSFORMATION RELATIONS
A27!A,
‘11 ‘12
‘12
‘n
‘16’ ‘26
‘16
‘26
‘66
- A;6 -A1866
A16A26
1 =
x
(30.A10)
I
-
A17!26
-
-
- A:2 There are several useful transformation rela- LA17!26 - A27!16 tions (prime refers to the off-axis coordinates 1,2,6 and no prime refers to the on-axis coor- where the determinant is IAl = (A,,A, - A:,)A6, dinates x, y, s as in Fig. 30.3). + 2Al.$26A16 - A,,A;, - A2,A:,. The flexural stiffness matrix [D] can be inverted similarly to {a’}= Ullal (30.A1) obtain the compliance matrix [d].
(4 = Ul-’la’}
(30.A2) 30.8 APPENDIX B: MORE DETAILS ON THE
QUADRATIC FAILURE CRITERION5
{&’I
(30.A4) 30.8.1 DETERMINING THE COEFFICIENTS
=
(Q’} = UIIQIVI where m = cos 6, n = sin 6 .
ul=
m2 n2 -mn
n2 2mn m2 -2mn mn m2--n2
m2 n2 -2mn n2 m2 2mn mn -mnm2-n2
(30.A5)
A useful exercise is to review how the F coefficients in the quadratic failure criterion
F,,a: + 2Fqaxay+ FyYai+ FS&
+ Fxax+ Fyay= 1 (30.B1)
(30.A6) are mathematically determined from the ply strength properties. During a uniaxial tension test of a 0 degree unidirectional laminate along the longitudinal axis (pulling in the direction of the fibers), ax = X and ay = os = 0. (30.A7) Substituting into eqn (3O.B1), the failure criterion reduces to
FxxX+ FIX = 1
(30.B2)
708 Laminate design
During a uniaxial compression test of a 0 degree unidirectional laminate, a, = X' and a = as = 0 and the failure criterion in eqn (30.Blf reduces to
familiar from isotropic materials. The most common form of the von Mises criterion is probably the equation written in terms of the principal stresses in three dimensions7.
FxxX'-FxX = 1
(a,- a,)' + (a, - a,), + (a,- a,)*= %2yield
(30.B3)
Solving the two equations given by eqns (30.B2) and (30.B3) for two unknowns, F,, and F,, yields 1 Fxx
=
xx'
1
1 F =--x X'
(30.M)
(30.B7)
For plane stress a, = 0, eqn (30.B7) becomes
a: - ala2+ a; = a*,,d
(30.B8)
Instead of computing the principal stresses, the von Mises criterion in eqn (30.B8) can also be written in terms of a general stress state that includes a shear stress contribution
Similarly for the transverse direction, a 90 - aXay+ 0; + 30: = (30.B9) degree unidirectional laminate can be experimentally evaulated under uniaxial tension (ay Now, for comparison, the quadratic failure cri= Y, a, = as =0) and compression (ay= Y', a, = terion is repeated here as = 0). FxXa: + 2Fya,uy + FW$ + Fssa:+ Fp, + Fyay= 1 Substituting each test situation into the quadratic failure criterion and solving the two (30.B10) equations, the resulting two unknown coeffiwhere cients, Fw and Fy,would be 1 1 1 F = F = - (30.Bll) F =(30.B5) YY Yy' xx XX' YY Yy' 1
1
~
1
1
Y
Y'
F =--!I
For a shear test, a, = S and the resulting coefficient is 1 (30.B6) F ss = -s2 The sixth coefficient, F ,must be determined empirically from biaxiar tests such that
FXY= F*&&) To have a closed failure envelope, Fey must be between -1 and 1. 30.8.2 RECOVERING VON MISES CRITERION
1
Fss = 7 Fy=F*yd(Fx,FYY) S 1 1 1 1 F =--F =--X X ' Y Y Y'
We can recover the von Mises failure criterion for isotropic materials when F*y = -1/2 and by setting the strengths to be the same in all directions X = X' = Y = Y',
S = X/d3 (30.B12)
where the 43 factor is the result of the von Mises invariant. Substituting expressions in eqn (30.B12) into those of eqn (30.Bll) and then rewriting eqn (30.B10) results in 0;
+ai
- O ~ O ~+
It may be useful to review the quadratic failure criterion and make it less of a mathematical which matches eqn (30.B9). abstraction by comparing it with what is
= x2
(30.B13)
DESIGN OF STRUCTURE WITH COMPOSITES 31 F.J. Schwan
to avoid confusion in this chapter. The term 'material properties' is used generally to Composite materials are finding more uses include extensional and compressive moduli each year across a wide spechum of applica- of elasticity, Poisson ratios, inplane shear modtions. Designers of structural components, in ulus, coefficients of thermal expansion and particular, continue to find new applications for coefficients of moisture expansion. 'Design reinforced composite materials. The most suc- allowables' refer to material strengths, specificessful structural applications are those where cally tension, compression, inplane shear, the innate advantages of reinforced materials interlaminar shear and bearing. This distinccan be translated into performance advantages tion between material properties and strengths for the manufactured part. These advantages is established because of the significant differinclude stiffness-to-density ratios, strength-to- ences which exist in our ability to predict and density ratios, low thermal expansion measure each group, and because of the differcharacteristics, and occasionally others, such as ent treatment which properties and strengths resistance to specific environments, thermal receive in all phases of the design process. conductivity, and fatigue characteristics. This section focuses on the design of com31.2 DESIGN PROCESS ponents determined by mechanical performance requirements such as stiffness, A general comment concerning the contents of durability, or strength. The importance of this section, and the chapter in general, is that establishing and articulating design require- the perspective represented throughout the ments cannot be overstated. The extent to chapter is that of a designer within a large which design requirements are isolated and organization, operating as part of a large quantified determines the degree of certainty design team, a situation typical of aerospace associated with trade studies and material projects. However, the process is the same for selection decisions. Unsuccessful applications all sizes of projects involved with design of are quite often the result of unclear or poorly composite structure. The primary differences are in the formality associated with each step in defined design requirements. Two terms which will be used extensively the process. In large organizations, extensive throughout this chapter are 'material proper- documentation tends to occur at each step, ties', and 'design allowables'. Since there are while in one-man teams little informal, and no universally accepted definitions of each of usually no formal documentation is required. these terms, it is necessary to adopt definitions The process itself, however, is the same. A typical conceptualization of the design process is presented in Fig. 31.1. While the Handbook of Composites. Edited by S.T. Peters. Published process is generically the same for all engineerin 1998 by Chapman & Hall, London. ISBN 0 412 54020 7 31.1 INTRODUCTION
710 Design of structure with composites Peliminary desian
Preliminary material selection
Member sizing - laminate design -joint design -tooling design
Detailed desian
Material specs
-
Manufacturing development
Process specs
7'
Design verification
Fig. 31.1 The design process.
Engineering development tests
- materals - critical elements - inspection techniques
I
- analyses - prototypes - testing - non-destructive evaluation
ing materials, the content of each box in the chart varies with many factors, including the materials of construction. The intent of this discussion is to provide some guidelines on what information is required to make key decisions and eventually converge on a design when composite materials are under consideration. The complex series of decisions which we refer to as the design process is commonly considered to comprise two phases: preliminary design and detailed design. Each phase requires a different level of requirements definition and application. Figure 31.1 indicates the information and data required to proceed in each phase. This data typically includes many kinds of information about the materials of construction and the fabrication processes to be considered. Design characteristics for each material are an important subset of this information. The objective of the preliminary design process is to develop a design concept to a level that includes configuration, materials, and preliminary assessments of performance, cost and manufacturing approaches. To support these determinations, a set of material characteristics data is required. As an example, key design requirements and a sample of what data is required to support preliminary design
Table 31.1 Design requirements for an aircraft sta-
bilizer ~~~~~~~~
Type of requirement Geometric constraints Production rate
Description Maximum length Maximum root and tip dimensions Number of units per month
Maximum weight; $ per kg of weight saved Minimum service life; + $ per additional hour of service life Environmental Service temperature extremes Chemicals - cleaning solvents, fuels, oils - concentrations - exposure periods
Cost/ benefits
Mechanical
Stiffness - maximum tip deflections - aeroelastic performance
Strength - imposed strains acceptable under design loads with factors -
of safety included fatigue life derived from required service life
Design process 711 Table 31.2 Preliminarymaterials data for aircraft stabilizer study
Type of data
Form of data
Available material forms
Prepreg -tape widths - thicknesses Pultruded forms - shapes Resin systems Specification details - fiber volume - void content - density - etc. Material availability
Manufacturing characteristics
Handling characteristics of material in automated tape laying process Cure characteristics - gel time - resin flow Prepreg storage conditions
Cost data
Material price Material development tests Manufacturing hours Tooling costs
Environmental resistance
Service temperature limits Moisture absorption characteristics Chemical resistance
Mechanical characteristics
Material properties
- extensional and shear modulus Allowable strains - tension - compression - shear Fatigue data - strain against number of cycles to failure of an aircraft component are shown in Tables The data required to support preliminary 31.1 and 31.2 respectively. The component design investigations include stiffness, denselected for this example is a vertical stabilizer. sity, strengths, fatigue characteristics, data at The design requirements define maximum tol- elevated and low temperature, forms in which erable deflections, imposed loads, a weight the material is available and associated matergoal, production rate, cost goal, service life ial prices. In addition, some experience with and operating temperature extremes. In real- the material in the fabrication process selected ity, the set of preliminary design requirements is required to support decisions concerning would be much larger including, perhaps, production rates and costs. Generally, reliable maintenance and inspection characteristics, published data can be used for preliminary descriptions of qualification tests, center of design. Only when no data is available is it gravity characteristics, and numerous others. necessary to generate data in this phase of the design.
712 Design of structure with composites The sequence of decisions in this design case would typically be to make a preliminary decision on a manufacturing technique, followed by selection of one or more materials based on cost and availability. Once this has been done, the component is 'sized' for stiffness, strength and service life in each candidate material. Material properties and strengths are used for these calculations. Material selection may depend on weight minimization or life cycle cost minimization or a combination of the two criteria. The result of this process is a material selection and a preliminary design. This 'design' possesses enough detail to support confidence that (1)the material selection is justified, and (2) that the design concept is feasible in this material. The set of mechanical properties and strengths was, by no means, a complete or final set of design values, but rather contains those characteristics needed to support preliminary design calculations. These values are generally the 'best available' from material supplier data, literature publications, data from other projects and data from material specifications. While the dominant design goal of aerospace designers is usually to reduce weight, the driving requirements vary widely from one product to another. Table 31.3 provides a synopsis of the diversity of these requirements across defense and commercial industries. This table indicates that requirements are numerous and, therefore, the derived design objectives are diverse and numerous. Payoffs are achieved in the form of weight savings or greater service life. Sometimes, as in the case of carbon-carbon rocket nozzles, the use of composite materials results in a new level of performance which is difficult to quantify due to a lack of economic alternatives. Table 31.3 shows a very large range in quantified payoffs. Each of these products reflect a history of successive applications and successive generations of the same application. With each successive development effort, the industry has become better at understanding and defining requirements, material characteristics, and payoffs. Since composite
materials incur increased engineering development, testing, and increased raw material costs, data from previous applications is quite valuable in performing cost trades and making preliminary design decisions. Table 31.3 suggests the ever-increasing scope of structural applications for reinforced composites. The requirements identified in this table are as diverse and varied as the components. The result of the wide range of requirements imposed upon composite materials, compounded by the ingenuity of material developers, has led to an overwhelming array of constituent materials. Many of these materials are very specific systems with a narrow range of applications. Some materials are suited to numerous uses. This multiplicity of materials and features has produced a plethora of design data. In order to make material comparisons and selections on a rational basis, the designer needs to establish a clear set of requirements for the material. These requirements are sometimes captured in a document which governs material procurement. The aerospace industry typically refers to these documents as material specifications. The situation in which the structural designer usually finds himself at the initiation of a design includes a broad and usually incomplete set of requirements and a mountain of literature on a vast array of materials. A design checklist, such as that contained in Table 31.4, may be of some help in getting started in this situation. This checklist offers a systematic approach to determining what information is needed at each step and suggests which design team members need to be involved at various points in the process. The structural design team will employ a process which resembles this one in order to develop a preliminary design which specifies configuration, materials, manufacturing process, and establishes a basis of confidence to proceed with final design. This basis comprises having answers or a plan to develop answers to all key questions.
c
a
Y a,
8
3
4 .3
Design process 713
714 Design of structure with coinposites
Table 31.4 Design checklist for composite structures What do 1 need to proceed with design of a structural component in composites? 1. A clear design objective
What am I trying to achieve?
2. Some quantitative measure of what that improvement is worth
3. A manufacturing approach
Are the tooling concepts feasible? Are the fabrication processes well defined? Do I need tool try units?
4. Material properties and allowables
Which characteristics are critical? How will I establish and verify these?
5. Material procurement
Will materials be procured to specifications? Do all specifications already exist? Will materials be available in suitable time and quantities? Is material development or tailoring necessary?
6. Joint concepts
How is the component to be attached to adjacent parts?
7. Design verification
Do I need special tests because I am using composites e.g. exposure to temperature or moisture? What kind of prototypes do I need? What tests will I subject the component to?
8. Inspection
What am I looking for? What are my accept/reject criteria? What techniques will I use?
Fig. 31.2 Design team.
Preliminary design 715 Figure 31.2 shows the makeup of a typical design team. For designs with composite materials, the team may be larger than design teams using other engineering materials. The focus of the team is the individual with design responsibility. Essential members are representatives of manufacturing, tooling, and materials procurement. Others might include materials suppliers, and specialists in testing and analysis. In some product development activities, the team includes subcontract managers or material procurement specialists who interact both with other team members and with other companies who also need to be considered as team members. More important than the list itself, is the early and continued involvement of all members. Interaction among these team members produces the design. As shown in Fig. 31.1, preliminary design begins with requirements definition and produces a design complete to the point of materials identification, fabrication approaches, size and shape, and preliminary lay-ups and joint designs. The development of this information requires the constant input and review of all key team members. One feature of the design team which is unusual is the inclusion of the material supplier in the design team. There are several reasons for doing this. Sometimes, existing materials do not 'fit the bill'. In these instances, further material development or modification may become part of the design process. The material specification can be looked at as a blueprint for the material supplier to use in design of the desired material. It is important to recognize the material supplier as a member of the design team so that his inputs and development efforts can be integrated into the design. Otherwise, some beneficial options may be excluded, both in terms of existing materials and near term material development efforts which may be of value to the project. 31.3 PRELIMINARY DESIGN
The discussions which follow concern the portion of the design process characterized as
preliminary design. The primary assumption made in this section is that the materials of construction are laminated composites with continuous fiber reinforcement. These materials are typical of structural applications. 31.3.1 MATERIALS
The subject of material design data, and the closely related subjects of material characterization and specifications, assume greater importance for composites than for conventional materials. There are several primary reasons for this: lack of a single, comprehensive design database or source; greater variability in measured values from one material to another and from one lot of the same material to another; differences between predicted and measured component performance introduced by variables associated with 'workmanship' in the manufacture of the part. The most basic reason, however, is the simple fact that the engineering development process encompasses design of both the material and the structure. For these reasons, the designer of composite structure needs to focus more attention, and usually more resources, on the subject of materials, than his counterpart using only conventional materials. Material data is required in order to make key design decisions, first in the preliminary design phase and later in detailed design. The type and extent of the data depends on the application, but, typically, data on the stiffnesses, strengths, and densities of candidate materials is needed. Figure 31.3 shows relative ranges of specific tensile modulus for composites with various fiber reinforcements. Specific modulus is a term which refers to the ratio of composite modulus to composite density. Specificmodulus is a measure of stiffness per kilogram of material, and is commonly used as an aid in material comparisons. Figure 31.4 shows the ranges of material costs for these same families of materials. Both of these figures were constructed considering the entire range of
716 Design of structure with composites
300.00 T cn
250.00 --
-aa 200.00 -$g
-‘8 3 150.00 -8
=a En u s2. 100.00 -z u 8
2
50.00 --
0.00
I
EGLASS FIEF!
ARAMID FlBER
,
I I
I
I
T300 GRAPHITE FER
I
M GRAPHKE
I I-M GRAPHITE
Material
Fig. 31.3 Specific tensile modulus of various composite materials. properties available from each specific material. These ranges are the result of including different forms of the material and the entire range of laminate values, including quasiisotropic values on one end and unidirectional values on the other. Figure 31.4 suggests some general conclusions. First, there are three more or less distinct regions of composites. The first group is glassreinforced composites. The second is intermediate modulus graphite and aramid fiber-reinforced composites. The third group includes high and ultra high modulus graphite fiber composites. Each of these types is suitable for a distinct range of applications. A comparison of where the common metals fall on this figure provides a pretty good explanation of why aramid and graphite have replaced metals in applications where the market can ’pay for performance’. Glass has replaced metals in those applications where some performance can be sacrificed for lower cost, increased environmental resistance or longer life. The second conclusion suggested by Fig 31.4 is that there is extensive overlap in the price of aramid and intermediate modulus graphite fiber composites. Selection from this group of
materials is usually made on the basis of performance. Figure 31.4 considers only material cost, which in some applications, is the major portion of product cost. In other applications, material cost is completely overwhelmed by manufacturing costs or other life-cycle costs. Material cost is typically about 30% of component cost in aircraft parts, less than 10% in satellite components, and generally higher than 30% in most commercial applications. It is therefore critical that product cost analysis include all important cost contributors. Generally, ’you get what you pay for’ or, more precisely, ’you must pay for what you need.’ However, the range of properties and costs within a family of materials is an indication of the latitude which a structural designer has when employing reinforced composites. The upper end of each range of specific modulus in Fig. 31.3 is established by values for a unidirectional laminate. While this is one equitable basis for comparing reinforced materials to each other, it is not a good basis to use for design calculations because so many requirements drive the design towards multidirectional reinforcement. Foremost among these is the need for properties in more than
Preliminary design 717 1000 900 --
800 --
P
'0°-600--
W HMSYMBOUZES HIGH MODULUS FIBERS INCLUDING GY70, M60J, PlOO & P120
500.-
IMSYMBOUZESINERMEDIATEMODULUS FIBERS INCLUDING T40, T50, T650 AND M40J
4
EI-
400-
9
300.200 -100 --
Fig. 31.4 Costs of composite materials.
3
I-
140
120
100
h
0 u)
3
80
#
5
t
s
60
40
20
0
Fig. 31.5 Mechanical properties of M laminates - T300/epoxy (0.60 fiber volumes).
718 Design of structure with composites one direction. Figure 31.5 shows the relationship of tensile modulus in x and y directions and inplane shear modulus (xy) for a varying lay-up angle, 13.The same type of relationship exists for strengths and thermal expansion coefficients. A design which maximizes modulus in one direction at the same time minimizes modulus in the direction transverse to this and minimizes inplane shear modulus as well. Strength in the transverse direction and inplane shear strength will be low in magnitude also. These low properties and strengths in the 'secondary directions' will severely limit the ability of the laminate to resist load and deformation in these directions. Highly directional laminates are therefore not appropriate for structures which act as plates or shells, and for beams with significant secondary loads. Highly directional laminates also present severe design constraints in joint regions where loads are multi-directional in nature. Multi-directional reinforcement offers a more robust design because strengths and properties are dominated by fiber properties in all inplane directions. Furthermore, laminates with multi-directional reinforcement offer more desirable cure characteristics and usually avoid high levels of residual stress, internal damage or warpage in thin parts, induced by cooldown from cure temperatures. Successful instances of unidirectionally reinforced parts are nearly non-existent. Table 31.5 provides a typical list of materials data required for structural design in rein-
forced composite materials. This table shows mechanical properties and strengths for four fiber-reinforced materials. These material properties and strengths are the basic data from which laminate values are calculated. As explained in Chapter 30, lamination theory simply uses layer values to determine engineering constants and to estimate strengths for any desired laminate. Materials testing is therefore conducted primarily at the layer level. Laminate test coupons are employed to confirm predictions and 'workmanship'. The exceptions to this statement are strengths which are laminate-specific, such as bearing strength and interlaminar shear strength. However, these are values which are important to only specific regions of the component, and generally do not enter into material selection decisions. For preliminary design, laminate values are typically predicted from available ply-level values. Design calculations and trades are conducted with these predicted laminate values. Later, in detailed design, predictions are confirmed with selective tests of design laminates. Carpet plots
The infinite number of combinations of ply thickness, ply angle and material reinforcement is both a curse and a blessing in preliminary design. Once the design requirements have been identified and quantified, the question of how to meet these requirements with any given material presents itself. An
Table 31.5 Typical materials data for preliminary design ~~
Material
~~
Mechanical properties Fiber volume Density Modulus fraction (dcrn3) ,Tong, Trans, (GPa) (GPa)
T300 Fabric M60J Tape PlOO Tape Aramid fabric
0.6 0.6 0.6 0.6
1.55 1.69 1.74 1.36
68.9 248 414 37.2
~
Strengths
68.9 5.52 4.14 37.2
Tensile
Compressive
Long. (MPa)
Trans. (MPa)
Long. (MPa)
Trans. (MPa)
620 993 827 827
620 27.6 27.6 827
414 496 414 138
414 172 138 138
Prelimina y design 719 extremely useful tool in addressing this question is the simple carpet plot. This graphical representation provides a way to present the entire range of properties, or strengths, available from a particular family of laminates. Figure 31.6 shows a carpet plot of tensile modulus for laminates comprised of layers of T300-epoxy with orientations of 0,45 and 90". The plot contains all possible proportions of the three directions, and is read as the legend indicates. This type of plot treats laminates as homogeneous combinations of various percentages of layers. Carpet plots do not recognize lay-up sequence. Generally, this presents no limitation, but for laminates composed of very few plies, or laminates employing significantly unbalanced stacking sequences, the assumption can lead to incorrect strengths. Carpet plots of material properties are constructed directly from a table of values for many specific laminates. Values of moduli, Poisson ratios, and expansion coefficients are calculated from lamination theory for a series of laminates with varying percentages of rein-
forcement in each of three specified directions. While 0, 45 and 90" plies are most commonly used, the three angles can be any selected values. Another common combination is 0/30/60. Typically, values are calculated using a difference of 10%between successive points, which requires about 1000 data points per set of curves. For this reason, creation of carpet plots is done with the aid of computers. Development of carpet plots of strength values is less straightforward. The primary complication is the definition of failure of the laminate. An extensive discussion of laminate failure modes is contained in Chapter 30. Calculation of laminate strength involves consideration of progressive failure of each layer and construction of strength envelopes for combined states of loading. Construction of a carpet plot requires an approximation of ply strength which is adequate for preliminary trade studies and initial member sizing. The recommended approach is to define laminate strength as the product of allowable fiber strain and the laminate modulus in the appropriate direction. This calculation assumes that
140
120 -0- 0% 0 LAYERS
++
10% 0 LAYERS 0 LAYERS -C-20%
4 3 0 % OLAYERS +40% OLAYERS 43- 50% 0 LAYERS +60% OLAYERS
4 7 0 % 0 LAYERS +80% OLAYERS +90%
0 LAYERS
+100% 0 LAYERS 0
I
1 0
10
20
30
40
50
60
PERCENTAGE4M)EGREE LAYERS
Fig 31.6 T300/epoxy x-directionmodulus.
70
80
90
100
720 Design of structure with composites the laminate contains fibers oriented in at least two directions, and further ignores the Poisson effect in the calculation of strength. While this approximation tends to overpredict strength, it has proven to be suitable for preliminary laminate selection in strength critical structure. Using this definition of strength, carpet plots can be directly constructed from allowable tensile and compressive strain values, and predicted modulus values. Carpet plots provide a quick method for selecting a candidate laminate for the preliminary design process. For example, let us assume that design calculations indicate that a modulus of about 90 GPa is required. The carpet plot of Fig. 31.6 offers a large number of laminates which meet this single requirement. One is a laminate comprised of about 55% 0" layers, 35% 45" layers and 10%90" layers. If tensile strength and inplane shear stiffness were important to the design, it would be necessary to consider the values associated with the selected laminate. Related carpet plots provide ranges of tensile strength (Fig. 31.7) and shear modulus (Fig. 31.8). These are about
427 MPa and 15.5 GPa, respectively. If this set of properties were not suitable, the three carpet plots would be used to iterate to an optimum set of values and an accompanying laminate. The use of carpet plots can greatly facilitate the laminate selection process in the early stages of design. It should be noted that these plots are simply visual representations of the results of lamination theory calculations, using ply level input values. Carpet plots of all inplane material properties: elastic moduli, Poisson ratios, thermal expansion coefficients and moisture expansion coefficients can be developed from the appropriate equations.
Design values Appropriate values for properties and strengths are required in both of the primary design phases - preliminary and detailed. For preliminary design, values are typically acquired from existing sources. These can include data published or provided by material suppliers in the form of brochures, design handbooks, data sheets, data published in
T
'0°
-D- 0% 0 LAYERS
2- 600 &w 500 A
X- 10% 0 LAYERS
+20% OLAYERS -A- 30% 0 LAYERS
I
a
b 400 w d
v)
300
+40%
0 LAYERS
+50% +60% -A-70%
0 LAYERS 0 LAYERS 0 LAYERS 0 LAYERS 0 LAYERS 0 LAYERS
+EO%
t-
+90%
+loo%
200
100
0
1
0
10
20
30
40
50
60
PERCEMAGE W E G R E E LAYERS
Fig. 31.7 TSOO/epoxy x-direction strength.
70
80
90
100
Preliminay design 721 40
T
t
0
0 7
0 01
0
r
n
0
d
0
0
0
In
(D
b
0 00
I
:
g F
PERCENTAGE&DEGREE LAYERS
Fig. 31.8 T300/epoxy x-direction shear modulus.
professional society proceedings or publications, data from textbooks such as this one, or data developed on previous projects within the organization. An abbreviated list of sources of preliminary design data is presented at the end of this chapter. As might be expected, this available data comes in various forms, so comparison of data from different sources may require some judgment and mathematical manipulation. Fiber volume fraction usually varies from one source to another. The data is sometimes normalized to a 'standard' value, typically about 60%. This normalization process is rarely explained in the literature, but most often consists of simply multiplying measured values by the ratio of 0.60 to whatever fiber volume was measured. The values most frequently normalized include tensile and compressive modulus, and tensile and compressive strength. The validity of normalizing compressive strength is debatable because compression strength of reinforced composites is dependent, to some extent, on the matrix. The validity of normalizing values other than the four mentioned above is dubious.
Test methods also tend to vary from one laboratory to another. In general, it is best to be wary of specimens other than the standard ones. Standard coupons and test methods are defined by the American Society for Testing and Materials (ASTM), and other organizations such as Suppliers of Advanced Composite Materials Association (SACMA). The primary value of employing standard coupons and tests is the confidence established with time and existing databases. While each standard coupon may possess some shortcoming, it does offer consistency and a basis of comparison with existing data. Fiber volume and coupon configuration are only two variants present among published data. Others include the resin system, material form, and method of fabricating panels from which specimens are taken. These variants underscore the reasons why existing data can be useful for preliminary design, but almost always needs to be augmented or replaced by data specific to the design application for final design. Two different approaches are employed in the development of design values. The first of these is to buy existing
722 Design of structure with composites commercial grade composite material, characterize it, and design to those values. The second is to include the material supplier in the process by jointly developing a specification to which the delivered material must comply. While the second approach may increase material costs, it guarantees the suitability of the material and usually results in the use of higher or more specific design values. A material specification typically defines acceptable values of all critical characteristicsof the material to be purchased. The specification may be a simple one page document speclfylng reinforcement, resin system, and limiting values of fiber volume, or a multiple page document including minimum values of modulus, strengths, cure characteristics, and other items. Material procurement cost will tend to follow the number and stringency of the criteria specified, since the material supplier will need more quality assurance activity before delivery and may need to perform additional material development.The more demanding a specification becomes, the greater the need to identify and work with one (or more) material suppliers in its development. One reason for this is simply to ensure the development of a specification which can be met by at least one supplier. A good specification needs to recognize the normal variation in material characteristics from one manufactured lot of material to another. Data illustrating this lot-to-lot variation will be shown and discussed in a later section of this chapter. 31.3.2 DESIGN EXAMPLE
Key aspects of the preliminary design process have been described in this chapter. An example, consisting of the design of a composite beam, is included here to illustrate the methodology typically employed in preliminary design, and to provide a demonstration of a trade study performed for the purpose of material selection. The structural element selected is a beam of constant cross section, fixed against
displacement and rotation at both ends, and loaded by a uniform load along the entire length, I, as shown in Fig. 31.9. Design requirements are summarized below: w = 2.0 nVmm
A S h
I = 1500 mm 4
Fig. 31.9 Example: trade study of beam with a uni-
form load. 1. Maximum beam weight is not to exceed 5.0 kg or 3.33 g/mm excluding end fittings. 2. Maximum beam deflection (at midspan) must be less than 2.5 mm under the load shown in Fig. 31.9. 3. Maximum stresses in the beam must not exceed allowables, including a factor of safety of 2.0. 4. The cross section of the beam must be closed and rectangular as shown. Four materials will be considered for this application: aluminum (7075 alloy), Eglass/ epoxy composite, T300 graphite / epoxy composite, and M60J graphite/epoxy composite. Formulas for deflection, maximum compressive stress, and beam weight, in pounds per inch, are presented below: Deflection: d = (w14/384EI) where €I = (Ebth2)/2 Max. stress: f, = M c / I f, = ( M / W = (wL2/4hbt) Weight:
W = v(2h+2b)t < 3.33 g/mm
One additional constraint must be introduced to the design. Buckling of one side of the rectangular section is governed by elastic moduli and by the width-to-thickness ratio of that side. The crippling strength of a side can be sigruficantly lower than the laminate compressive strength or
Detailed design 723 inplane shear strength. In order to prevent compressive or shear buckling of the sides and flanges, a rule of thumb is to limit the side dimension, h, to no more than twenty times the thickness, and to limit the flange dimension, b, to no more than fifteen times the thickness. Using these relations, h and b are eliminated from the equations and replaced by multiples of t. The three equations are then solved by iteration for each of the four materials using properties and strengths shown in Table 31.6. The summary table presents a comparison of the calculated design thicknesses and corresponding beam weights for each of the four candidate materials. The deflection requirement proved to be the dominant requirement for all four materials, i.e. the thickness required to satisfy the deflection requirement was greater, in each case, than that required to limit imposed stresses to acceptable values. The estimated preliminary design weight of each design is less than 5 kg, as required. Several conclusions can be drawn from this comparison. First, the aluminum and E-glass designs are comparable on a weight basis due to the fact that the lower density of fiberglass offsets its lower stiffness. End fittings have been ignored, along with numerous other details, in this preliminary study, but it is worth noting that the additional weight introduced by metallic end fittings would probably cause the E-glass detailed design to be heavier than the aluminum design. The T3OO/epoxy design offers a significant weight savings over aluminum and E-glass, deriving primarily from its higher modulus-to-density ratio.
However, the cost of a beam fabricated in T300epoxy may still exceed the cost of an aluminum beam, despite the fact that less T300 material is required. The next calculation might be to estimate the fabricated costs, including materials, of the T300 and aluminum designs, to ascertain the cost differential. A final comparison can be made between T300 and M60J fiber reinforced beams. As presented in Table 31.6, the additional weight savings associated with an M60J beam is relatively small. The cost differential is going to be substantial because M60J prepreg is over $200 per kilogram while T300 prepreg is about $20 per kilogram. Only a substantial payoff in increased component life or overall system performance would justify the selection of M60J over T300. In a real design situation, the trade would probably be more complex. Perhaps, the load would be a cyclic load or an intermittent load, leading to consideration of fatigue allowables. Or, perhaps the natural frequency of the beam is another important design consideration in the structural system. The trade methodology remains the same, but additional characteristics would need to be considered. 31.4 DETAILED DESIGN
Many of the activities which constitute detailed design are extensions or iterations of activities already performed in preliminary design. However, a few topics require special attention. This section will focus on three areas specifically. These include a discussion of the development and use of material design data,
Table 31.6 Results of design trade study of composite beam
Material" Alum (7075) E-glass T300 M60J
Modulus (GPa) 68.9 34.5 99.3 248
Compressive strength (MPa)
Density (g/cm3)
482 475 572 379
2.68 1.93 1.50 1.69
Wall thickness
hm)
Weight (g/mm)
2.7 3.2 2.5 2.0
1.37 1.38 0.66 0.47
a Design values for each composite material are developed assuming a laminate with the following proportions: 65% 0" plies; 25% 45" plies; and 10%90' plies.
724 Design of structure with composites the approach to joining and the development of joining details, and the construction of drawings. 31.4.1 USE OF MATERIAL DESIGN DATA
The first of these three areas is an extension of the discussion of material data introduced in the section on preliminary design. The point was made in that section that material data relies heavily on ply-level values. A typical set of mechanical material data is shown in Table 31.7. The material of interest is T50 graphite fiber in an epoxy resin system. The material was ordered in significant quantity to an aerospace specification and was manufactured in ten lots. Table 31.7 is subdivided into material properties and strength values. The values shown as 'measured' are average for all specimens from all ten lots of material. The table also includes preliminary design values for the sake of comparison. This comparison shows excellent agreement between measured and preliminary design material properties. The largest differences are in coefficient of thermal expansion (CTE) values, and these are on the order of 20-25%. This is excellent agreement considering the difficulties and uncertainties associated with the
measurement of this property, and the expected variation in this characteristic. Differences in design values and measured strengths are larger. However, preliminary design values are all exceeded by average measured values, confirming the fact that strength critical regions are conservatively designed. There are some differences in values measured from one lot of material to another. The kind and amount of variation is described in more detail in Figs. 31.10 and 31.11. Specified 'acceptance values' are shown on each figure for reference. In this case, these acceptance values are lower than those used for preliminary design. The figures also show results of tests conducted by the material supplier and by the receiving company. Differences in these results provide some insight into the influence of test methods, facilities, and personnel. The first figure compares tensile modulus and the second tensile strength. Figure 31.10 indicates that, while values are different from lot to lot, all lots exceed the specification value for tensile modulus and are therefore acceptable. The measured ply-level data confirms the design values of tensile modulus. The measured values for tensile strength exhibit larger variations, but are also higher than the specified value. Once again, all measured values
Table 31.7 Mechanical design values for T50/epoxy tape: preliminary and measured layer properties and strengths Material characteristic Tensile modulus, E,,(GPa) Tensile modulus, E,,(GPa) Poisson ratio, v Inplane shear modulus, GI,, (GPa) Longitudinal CTE (x l P / " C ) Transverse CTE (x 1O4/'C) Tensile strength, F;, (MPa) Tensile strength, FZt,(MPa) Compressive strength, FIC,(MPa) Compressive strength, F;, (MPa) Inplane shear strength, F,,, (MPa) Subscript 1 denotes fiber direction. Subscript 2 denotes direction transverse to fiber. Subscript 12 denotes shear in the plane of the fiber.
Prelimina y design value 209 6.89 0.27 5.52 - 0.63 28.8 1034 24.1 758 110 62
Measured value 228.2 7.03 0.245 5.44 - 0.79 25.0 1136 28.06 821.2 171.6
Detailed design 725
300
T
iii 250
n
9
; v)
-
200
0 I W
LMSC DATA HEXCEL DATA
d 150 6
+LMSCSPEClFlCAl-ION VAI-UE
I-
0
1
2
3
4
5
6
7
8
9
10
LOT NUMBER
Fig. 31.10 Lot-to-lot variation in tensile modulus - T50/F384 unidirectional tape (manufactured by Hexcel).
1400
1
g 1200 n
3
I
6 1000 E
F
800
HEXCEL DATA
d
mLMSC DATA
u)
+LMSC SPEClFlCPiTlON VAI.LIE
z 600 +
W
0 1
2
3
4
5
6
7
8
9
10
LOT NUMBER
Fig. 31.11 Lot-to-lot variation in tensile strength - T50/F384 unidirectional tape (manufacturedby Hexcel).
726 Design of structure with composites exceed specification levels and all lots are acceptable. This data is typical of aerospace applications. The types and amounts of data will, of course, depend on the specific application. The major difference between this data and preliminary design data is that the data supporting final design will be specific to the material form, resin system, and fiber volume selected. Final design data is an extension of data available for preliminary design. It will generally focus on ply level values since this is the most cost-effective way to obtain data for all of the laminates which will be employed in final design of the part or assembly. These values now become an important part of the
database for establishing preliminary design values for the next application. Table 31.8 contains a further comparison of predicted and measured laminate values for T50-epoxy tape, developed in the same application. The values in this table correspond to specific design laminates. Three sets of values are shown for each of the design laminates the first developed from preliminary design layer values, the second from measured layer values, and the third directly from tests performed on laminate coupons. Using measured layer properties produces a better correspondence between predictions and measurements, but not an exact overlay. The primary difference is workmanship, i.e. changes introduced
Table 31.8 Comparison of measured and predicted material design values for T50/F584 graphite epoxy tape laminates
Design value
Laminate description (60/0/-60/0,)~
(45/45/0,)s
Prelimina y design
Prediction
129.5
141.1
148.9 (60)
51.4
55.6
0.303 19.8
Tensile modulus, Ex, (GPa) Tensile modulus, Ey' (GP4 Poisson ratio, v Inplane shear modulus, Gx ,(GPa) LongitudinaICTE (x lOd/OC) Transverse CTE
Measurement
Prelimina y design
Prediction
Measurement
115.4
125.2
129.6 (10)
51.1 (60)
28.8
30.8
26.8 (10)
0.303 21.2
0.289 (60)
0.762 29.3
0.776 31.8
0.777 (10)
-
- 0.58
- 0.49
- 0.41 (5)
- 0.67
-1.08
-1.17 (3)
- 0.29
+1.55
+2.02 (6)
- 0.01
+3.51
+4.10 (3)
641
703
693 (60)
565
578
781 (10)
324
341
170 (60)
255
284
116 (10)
469
507
623 (60)
414
446
605 (10)
255
278
165 (60)
255
235
128 (10)
165
134
-
214
209
-
-
(x lP/OC)
Tensile strength, F,',
Tensile strength, F t, (MPa) dbmpressive strength, F;, ( m a ) Compressive strength, F;, (MPa) Inplane shear strength, Fm, (MPa)
Number of test specimens shown in parentheses alongside measured values. Subscript x denotes the 0" direction. Subscript y denotes the 90"direction.
Detailed design 727 by handling and fabrication that simply are not accounted for by theory. These differences are larger for strengths than for properties, and account for the knockdown factors normally used in developing material design strengths. One final point to be made concerning these comparisons is that differences exist, but are not typically an inhibitor to designing with composites. These variations must be recognized in the design process, and can be handled with statistically-derived knockdown factors, and conservative but realistic design approaches. Key design values can also be controlled through development of intelligent material specifications. In most designs, the usual variations are easy to accommodate. The important point is that they need to be quantified and included in the design process. This leads to a discussion of the role of testing in the development of design allowables. It was previously pointed out that normally little or no testing is performed in support of preliminary design. In fact, the only testing which is essential is that performed to determine critical values which are just not available. In the process of detailed design, however, the picture changes. Once a material is selected and a design acquires some maturity, the values critical to establishing sufficient confidence in the final design can be identified and addressed through test. The emphasis is usually placed upon layer level coupons for two reasons. The first is that the information obtained from unidirectional, standard coupons is the most direct information available about the material. Failure modes are predictable and repeatable, and results can be directly converted to allowables. An additional benefit deriving from the use of unidirectional coupons is that direct comparisons to other materials and different lots of the same material can easily be made. The second reason for allocating the majority of testing resources to ply-level testing is that this data can be used for prediction of a wide range of laminate properties and
allowables, and so provides maximum leverage in terms of applying testing resources to development of design confidence. A less extensive series of laminate tests may be required in addition to the unidirectional coupon testing. The purpose of these tests would be to provide some selective confirmation of key design properties and strengths for comparison with values predicted from measured layer values. In addition, some very specific strength characteristics, such as bearing strength, are specific to laminates, and cannot be reliably predicted from layer values. If this allowable is critical to the design, tests of specific design laminates may be necessary. 31.4.2 JOINING
Joining of composites normally presents some challenges and tends to influence the configuration of the part. The two recognized methods for joining composite structure to other composite or to metallic parts are adhesive bonding and mechanical fastening. The advantages and disadvantages of each method will be discussed, followed by some guidelines for selection between the two. The discussion is not an exhaustive one by any means, but is intended to highlight key considerations which determine joint designs and provide some initial direction to the designer. Numerous factors need to be considered in the selection process. Figure 31.12 offers a list of the most common requirements. It should be kept in mind that some of these 'requirements' may become design variables in the course of the process. Geometry of the members being joined, for instance, could be altered locally to facilitate joint design. Reliability potentially includes an array of requirements, one of which is the implications of joint failure on system performance. Is it catastrophic or benign? Is it easily detectable prior to failure or insidious? Determination of this list of requirements will lead directly to the formulation of key
728 Design of structure with composites
Joint design criteria
1 Temperature range Moisture limits
-Static strength -Durability (fatigue)
How important is weight? How critical is cost?
i , t Reliability
Service life Implications of fai Iure
Fig. 31.12 Joint design. design questions possibly including, but usually not limited to the following: Will the joint be disassembled and reassembled? What is the nature and magnitude of the loads? Are the loads cyclic or static or a combination? What are the operating temperature and moisture ranges? What are the cost limitations? Some of the major advantages of bolted joints and bonded joints are shown in Fig. 31.13. High tolerance to repeated loads, good resistance to most environments, ease of inspection, and high reliability are primary advantages offered by mechanical fasteners in composite structure. Fewer pieces, lower weight, good load distribution, and lower cost are advantages offered by adhesive joints. The major considerations in design of a bonded joint can be grouped into five categories. These include joint strength, environmental resistance, joint geometry,
Mechanically fastened
- fatigue life - insensitive to temperature - ease of inspection - reliability
1 Adhesively bonded - no local reinforcement (load spreading) - fewer parts - lower cost - corrosion resistance
Fig. 31.13 Advantages of bolted and bonded joints.
selection of the adhesive system, and processing. Extensive information on each of these subjects is available in other sections of this text. The following design procedure is offered as a facilitator to the design process. Generally, component configurations are predetermined by other design requirements,
Detailed design 729 and so the joint geometry is established prior other sections, and so this discussion is aimed to finalization of the adhesive joint design. at providing a brief summary of the process. The process begins with the determination Factors including how well the mating parts will 'fit up' to each other after allowances for of a configuration for the joint. Single lap joints all specified dimensional tolerances, and the are normally adequate for thin laminates (up quality of the adherend surfaces need to be to about 5 mm (0.2 in) in thickness). Fastener bending and initial bearing failure are primary recognized and defined. The first step in the process is to determine concerns. Double lap joints are better for cyclic a dimensional configuration which minimizes (fatigue) loads, and generally stronger. The next step is to select the type of fastener tensile and peel stresses in the bonded joint. Once this is accomplished, the next task is to required. Fastener selection usually raises select an adhesive system which best satisfies issues requiring decisions concerning laminate static strength, fatigue life, and environmental reinforcement, hole sizes and locations, requirements. This selection is sometimes the drilling, fastener installation, and inspection. result of a simple trade study among several Table 31.9 identifies issues and proven design candidates. The third step is the development approaches to each issue. The table reveals of formal or informal process specifications for that the complexity of designing bolted joints the joint. This specification will include details derives from two primary sources: (1) comfor surface preparation, curing the joint, and posite laminates cannot redistribute high local maintaining pressure during cure if necessary. loads by yielding and plasticity; (2) composThe joint strength is typically verified analyti- ites are more easily damaged by drilling and fastener installation than metals. cally or by structural test or both. Design of local reinforcement of the lamiSimilarly, a design process can be outlined for mechanical joints. The primary design con- nate to resist local stresses is an important step siderations for bolted joints include joint in the design of bolted joints. If reinforcement strength, fastener type, local reinforcement, is required, a proven approach is to increase joint configuration, holes, and preload. All of laminate thickness by the addition of plies these subjects are discussed extensively in placed at 345 and 90" to the primary load
Table 31.9 Design issues: fasteners in composite laminates
Issue
Solution approach
Drilling damage
1. Closely controlled manufacturing operations 2. Inspection of drilled holes
High local stresses
1. Larger fastener diameter 2. Insert (bushing) 3. Increased laminate thickness (locally)
Preload relaxation
1. Large fastener head 2. Washers (one or both sides) 3. Limit on installation torque
Countersunk head
1. Avoid, if possible 2. Increased laminate thickness (locally)
Damage induced by installation of blind fasteners 1. Specially-designedblind fasteners 2. Development of modified installation techniques and drive rivets 3. Verify joint strength with tests
730 Design of structure with composites direction. A quasi-isotropic laminate provides the best bearing strength in any continuous fiber polymer composite.
Local reinforcement As in metal structures, local reinforcement is generally required where any hole or cutout is placed in a structural part. Analogous to metal parts, reinforcement can be bonded or fastened to the structure. Reinforcement can often be incorporated into the design of the part in the form of additional plies extending across a limited region of the structure. Local reinforcement may also be required in the vicinity of joints, either bonded or bolted, and in locations where concentrated loads are introduced into the structure. The design of the reinforced region will derive primarily from the orientation and magnitudes of the stresses in the structure, and so no generalizations can be made concerning laminate design. However, some generalizations concerning transition in these regions and the behavior of laminates are appropriate. Laminates behave as a series of layers, each stressed in membrane action, and each capable of transferring stresses to adjacent layers through shear. When laminates bend, adjacent layers remain attached and experience states of strain which are similar. Therefore, in a reinforced region, such as that shown in Fig. 31.14, load is introduced into the reinforcing plies through interlaminar shear. Since the level of interlaminar shear stress required to induce delamination in any composite will not exceed 70-80MPa, a design must aim to maintain these stresses to acceptable levels. This is typically done through the gradual introduction of reinforcing plies, as shown in Fig. 31.14. This gradual transition precludes a sudden discontinuity in thickness which would produce sharp peaks in interlaminar shear stress. The selection of ply orientation is also a consideration. Transfer of load from ply to ply is facilitated by the minimization of differences
4 x 0.25
, w Scale: 118 _f
0.10 (reference).
f'
(1 .O) See view B view _ f ~ . 2(reference) ~
.- .K,
I I \
'--' A-A Scale 1:2
Fig. 31.14 Local reinforcement.
in the stiffnesses of the two plies. Therefore, load transfer is most effective between two plies with the same fiber angle and least efficient between two plies with a 90" difference in fiber angle. The use of 45" plies between 0 and 90" plies is a good compromise for reinforcement areas. Analysis methods are limited by underlying assumptions, and are most reliable for regions where the assumptions are simple and few. This is not the case for regions in which the laminate changes and the stress gradients are steep. Therefore, complete reliance on analysis in the regions of joints and local reinforcement is not prudent. In addition, these regions sometimes present manufacturing challenges which raise questions concerning the quality of the finished part. Therefore, testing is very often an integral part of engineering development and of the design verification process. 31.4.3 DRAWING CONSTRUCTION
A drawing of a component or assembly in composite materials has the same objective as this type of drawing in any material: furnishing a blueprint for manufacture of the item. However, some of the details and the manner in which information is presented on the
Detailed design drawing are unique. There are countless variations on the format of a drawing, so the discussion will be more or less limited to the content, with a specific format shown only for the purpose of providing an example. A drawing detailing a composite part needs to specify geometry, materials, sources of those materials, ply sequence for each unique laminate, details of transition regions, processing definition, post fabrication instructions for drilling, sawing, etc. and any inspection requirements. Other details, such as individual ply patterns, may be necessary for a specific part, but this list is suggested as one of general applicability. Dimensions are defined in the same manner as for metal parts with the exception of thicknesses. Thickness is really defined by laminate sequence meaning that the part will be as thick as the cured laminate dictates. A thickness dimension is normally shown as a reference dimension and provided for tooling design. It should be recognized that specifymg a thickness as a fixed dimension demands compliance to that dimension. Unless manufacturing ‘tool try’ units have been fabricated to determine cured laminate thickness with the particular manufacturing process specified, there is no guarantee that the thickness can be achieved with the number of plies and resin content specified. Dimensional tolerances are generally determined by tooling design and the amount of material shrinkage. Experience with the process and material is the only reliable guide in setting tolerances. Material definition can be accomplished in two ways. The first is to create a note or a
table defining the material: fiber reinforcement, resin system, material form, resin content, material source or sources, and surface treatments. The second is to invoke a document, such as a specification, which defines the material required. If the second approach is used, it may still be a good idea to provide a brief identification of the material on the face of the drawing to facilitate interpretation by analysts, tool designers, manufacturing and inspection personnel. Table 31.10 shows an example of material description employing both methods. Specifying a fabric requires the designer to be familiar with all of the variable characteristics of the material form - weave style, surface treatment, and end count. This is necessary because fabrics are available in numerous weave styles, with various surface treatments, and can be woven with any of several different size yarns. Each unique laminate in the part requires definition sufficient to allow a fabricator to layup the part. Figure 31.15 presents a typical example of the usage of a ply sequence table on the face of a drawing. The figure identifies the surface that ply number 1 is placed upon and then details the ply stacking sequence to be followed. It is desirable, but not always practical, that the first ply shown in the table be the first ply physically laid up on the tool. The ply table contains a column identifying the material of each ply since sometimes a single laminate contains plies of different materials. Areas of transition are handled as shown in Fig. 31.16. The two adjacent regions of ‘constant’ thickness are defined by appropriate tables as shown. The tables, however,
Table 31.10 Format for material definition on drawings
Material designation
731
Characteristics
Source
M1
Fabric prepreg, 120 style weave, T300 fiber, 10K bundle size, 934 resin system, sizing
Company name and address
M2
M40J tape
Specification number
732 Design of structure with composites Ply number
-b
1 2 3 4 5
6 7 8 9 10
Material designation
Fiber orientation
M1 M1 M2 M1 M1 M1 M1 M2 M1 M1
90 0 45 90 0 0 0 45 0 90
("1
Fig. 31.15 Ply sequence table. Material -iber angle (") 90 0 -45 45
M1 -45 45 0 90 10 11 12
M2 M2 M2 M1 M1 M2 M2 M1 M1 M2 M2 M1 M1 M2 M2 M2
90 0 0 45 -45 90 0 45 -45 0 90 -45 45 0 0 90
Fig. 31.16 Specifymg transition regions.
employ a common ply numbering sequence so that the continued and terminating plies can be distinguished. The termination details are specified on the face of the drawing as shown in Fig. 31.16. Processing instructions are supplied in one of two ways, either as notes, or by reference to a document such as a process specification. The second alternative is by far the most common because of the amount of information required to define the process. Process
specifications typically include definition of cure cycle (temperature and pressure against time), pre-cure instructions for vacuum bag, bleeder plies, resin preheating for pultrusion, tool preparation, removal of the part from the tool, in-process inspection requirements, and many other items. Post-fabrication operations may include drilling holes, sawing to length, removing edges by grinding, secondary bonding of items, painting, light sanding, and almost
Detailed design 733 anything else. These operations are normally specified on the drawing by means of notes. Notes need to identify items such as the type of drill bit and cutting speed.
The advantage of documenting each NDE activity is traceability. If a problem arises in development or production, NDE records can usually identify potential sources and trends. The disadvantage is cost. Each company performs its own cost/payoff determination, and 31.4.4 DESIGN VERIFICATION PROCESS develops its own verification processes from A few words about design verification are that determination. One point for the designer appropriate to a section dealing with compo- to keep in mind is the necessity and value of nent design. Design verification encompasses this verification for composite parts. Variability more than the nondestructive evaluation in raw materials, processes, and workmanship (NDE) procedures specified on the compo- can and will affect the final part. The designer nent drawing. Design verification is integral needs to be aware of all potential sources of to the entire design and manufacturing variation, so that he knows where to look when process. Figure 31.17 identifies verification final inspection reveals defects in his part. activities in each phase of the component Part inspection methods are usually identidevelopment process. Each of these activities fied on the drawing. These requirements ranges from a formal, documented step in actually form the accept/reject basis for the some industries, to undocumented, highly finished part, and can range from visual only integrated steps in other industries. The key to extensive ultrasound methods. Table 31.11 point is that verification is the part of the presents a summary of which defects are development process which ensures that the detectable with each of several established developed component fulfills all design NDE techniques. The most commonly used methods are X-ray and ultrasonic, particularly requirements.
Primary Activities
Development phase
Verification actions Development tests
Engineering development
Material Procurement
Material acceptance tests
Tool tries In-process inspection Fabrication
F) fabrication
Processing records
Debulk Bagging
Tag end coupon test
Trim
Witness panels
Bonding
Dimensional verification
Drilling holes Sawing
Non-destructive inspection
acceptance
Fig. 31.17 Component verification process.
Proof test / acoustic emissions
734 Design of structure with composites Table 31.11 Defects and nondestructive test methods
Defect or variable Unbond Delamination Undercure Fiber misalignment Damaged fibers Variation in resin Variation in thickness Variation in density Voids Porosity Fracture Contamination Moisture
X-ray
Gamma “Y
X X
X X
Ultrasonic
Sonic
Microwave Temperature Penetrant diferential X
X X
X
A
X X
X
X X X X X X
X X X X X X
X
C-scan. These methods require proper equipment, cumbersome setup in some instances, and time. X-ray and ultrasonic methods can detect a wide range of defects, and are sensitive in ranges where other methods are not. However, these methods have definite limits. In fact, no single method can detect all types of defects or variations. Moreover, it is important to understand, in advance, what a defect looks like. For this reason, it is common practice to perform calibration with intentional defects built into test panels. Table 31.11 is an oversimplification of an extensive and complicated subject. It is included here as a handy reference for the designer and a starting point for determining which techniques are appropriate for a given part. Inspection requirements will, of course, be related to the function of the part, and will be focused on the critical aspects of the design. In some companies, it is common practice to limit part inspection to simple dimensional and visual checks, and to rely on a ‘proof’ test to verify the structural adequacy of the component. Since inspection requirements tend to be unique to each component, it is usually necessary to make inspection requirements an integral part of the design drawing.
X X X
X
X X X
X
SOURCES OF MATERIALS DATA FOR PRELIMINARY DESIGN
Military Specifications
Mil-R-9300B
Resin, Epoxy, Low Pressure Laminating Mil-C-47257C Compound, Epoxy, Filament Winding Mil-R-24719 Resins, Vinyl Ester, Low Pressure Laminating Mil-Y-1140 Yam, Cord, Sleeving, Cloth and Tape, Glass Mil-R-60346 Roving Glass, Fibrous (for Filament Winding Applications) Mil-Y-83371 Yarns, Graphite, High Modulus, Continuous Filament Mil-Y-87125A Yam, Graphite, 1000/3000 Filaments Mil-F-87121A Fabric, Graphite Fiber Mil-T-29586/ 1 Thermosetting Polymer (Epoxy) Matrix, 350°F (177°C) Cure, Intermediate Modulus Carbon Fiber Reinforced Prepreg Tape
Sources of materials data for prelimina y design 735 Aerospace Material Specifications Glass Roving, Epoxy Resin AMS 3828 Preimpregnated, Type E Glass AMS 3892/8-84 Tow or Yarn, Carbon (Graphite) Fibers - For Structural Composites AMS 3894/11-83 Carbon (Graphite) Fiber Tape and Sheet, Epoxy Resin Impregnated Cloth - Type 'E' Glass, 'B' AMS 3822 Stage Epoxy Resin Impregnated, Style 181-75DE
Modern Plastics Encyclopedia, McGraw Hill, New York, yearly edition. Composite Design Encyclopedia, University of Delaware, Newark, 1984. Composite Materials Handbook, (ed. Me1 M. Schwartz,) McGraw Hill, New York, 1984. Periodicals
Modern Plastics Magazine, 1221 Sixth Avenue, New York, monthly periodical S A M P E Quarterly Journal
Federal
Journal of Advanced Composites
Mil-Handbook 17C, Polymer Matrix Composites, Volumes 1 and 3, Standardization Documents Order Desk, Building 4D, 700 Robbins Avenue, Philadelphia, PA 19111-5094
Material Suppliers
Professional Society Publications
Reinforced Plastics By Design PPG Industries, Fiberglass Reinforcements Market Series
Plastic Design Guide Owens/Coming Fiberglas Corporation, 1974
Society for the Advancement of Material and Process Engineering (SAMPE)- Conference Proceedings
Hexcel Design Guide
Society of Manufacturing Engineers (SME) Conference Proceedings
Data Sheets from material suppliers:
Handbooks
Handbook of Reinforced Plastics. (eds. S.S. Oleesky, J.G. Mohr). Van Nostrand Reinhold, New York.
ICIFiberite Materials Handbook ICI/Fiberite Hexcel Aerospace Products AMOCO Performance Products Nippon Oil Company, Ltd.
ANALYSIS METHODS
32
V V Vasiliev
32.1 INTRODUCTION
A detailed analysis of composite structures can provide information about the stress-strain state, strength, stiffness, stability and vibrational behaviour of structural elements. The analytic methods are those used in solid mechanics; the equations are modified to reflect the structural behaviour of composite structures. These equations are based on a system of assumptions that take into account the key features of a structure and establish the
EQUl L I BRI UM EQUATIONS
LOROS
I ENUIRONMENTRL RNO OPERRTI ONRL CONDITIONS r
I
STRAIN-DI SPLRCEMENT EOURTIONS
+
I
--+-b
appropriate model for the structural element under consideration. To start the analysis (Fig. 32.1), one must know the geometric configuration of the structure, material properties, environmental and operational conditions along with the applied loads. It must be noted that the available data on material properties and loading conditions affect accuracy, duration and cost of the analysis. As an illustration, consider a cylindrical shell loaded with a concentrated radial force. The solution for this problem can be obtained
RSSUMPTIONS A N 0 MOOELS
GOUERNING EQURTIONS
A
I
CONSTlTUlTlUE EQUATIONS
I
EXPERIMENTAL RNRLYSIS
I-
*
OUERRLL STRRINS RNO OISPLRCEMENTS
I INLRMINRR
STRENGTH
-
EXPERIMENTAL UERlFlCRTlON
Handbook of Composites.Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Fig. 32-1 Outline of analysis for composite tures.
StrUC-
Introduction 737 in the form of a Fourier series that has a very poor convergence at the loaded point. However, a closer inspection of the loading conditions often shows that the concentrated force is the resultant of some distributed load over an infinitesimal area; this distributed load can then be expressed by the series with a finite number of terms to achieve convergence. Operational requirements for the composite structures are also of great importance. For example, a filament wound composite shell of revolution can be a model for either a gas pressure vessel or a solid propellant rocket motor case for which, in contrast to the pressure vessel, time dependent material properties need not be considered. On the basis of this knowledge, assumptions are formulated for the model of the structural component under study. This model should consider only the key factors affecting the capability of the structure and ignore the secondary effects that complicate the analysis without significant improvement. The next step implies analytical formulation of the problem, i.e. development of a complete set of the governing equations with the pertinent boundary conditions. For structural analysis, the governing set of equations consists of equilibrium equations (or equations of motion), strain-displacement and the constitutive equations which are material dependent. The constitutive equations include stiffness coefficients which are determined by either of the following two methods: 0
0
the experimental method can not be used for unique structural designs. To minimize the degree of uncertainty, a second method is applied. According to this method, the stiffness coefficients are calculated from data that specify the composite material, ply co-ordinates, ply thickness, ply orientation and mechanical properties of individual plies. The latter, in principle, can be predicted by analytical or numerical methods of micro-mechanics of composite materials; these results, however, have more academic than practical value. Idealized micro-mechanical models, that approximate the real structure of the material, do not account for material porosity, non-uniform fiber distribution, possible variation of physical properties in each batch of resin, variation in strength of the resin/fiber interface and a number of other factors associated with the manufacturing process; in many instances, the above variations can not be described in formal analytic terms.
For important structures, usually, the mechanical ply-properties are experimentally determined by testing specifically fabricated plates, rings, or tubular unidirectionally reinforced specimens produced by the same process as the structure under study'. Once all of the pertinent data are together, the governing equations are solved by either the analytic or the numerical methods applied in solid mechanics. The results by either The first method is associated with direct method yield information on deformation, experimental analysis of test specimens cut displacements, stiffness and strength. To evalout of the structure or its excess length. This uate stiffness, the calculated displacements method is used mainly for structures in due to an applied loading can be used mass production. For simple laminates (e.g. directly. Evaluation of strength requires transfor cross-ply laminates), this method gives formation of the calculated global strains to rather accurate results while for the more ply strains in terms of principal material cocomplicated (e.g. including angle-ply lay- ordinates, determination of the ply stresses ers) laminates, separation of a specimen can and application of some strength criterion. cause a disruption in the material structure One must note that the numerous strength criand induce specific edge-effects with ques- teria, that have been developed to date, are no tionable mechanical properties. Obviously, more than analytical approximations of the
738 Analysis methods experimental results with a significant scatter. Also, the failure of a composite material is a rather complicated process that defies accurate theoretical description. Therefore, the strength of important structures should be always verified by testing either sub-scale or full-scale models. Finally, a thorough analysis of a composite structure, based on macrostructural models, effective stiffnesses and approximate displacement fields, does not preclude the possibility of a detailed analysis of inlaminar or interlaminar stresses. To accomplish this analysis, the ply under study is separated out of the structure, loaded with unknown contact forces, deformed in accordance with the known global strains of the structure and described by the proper microstructural model.
't
SI
Fig. 32.2 Element of a composite structure.
1. Equations of motion 32.2 GENERAL EQUATIONS FOR COMPOSITE STRUCTURES
Composite structural elements consist of a large number of plies with different angular orientation of fibers. The general equations, describing the behaviour of the composite elements, are based on the following assumptions: 0
0 0 0
the structural element consists of perfectly bonded anisotropic layers made out of linear elastic materials; the laminate thickness (h) is much smaller than the other dimensions; the radii of curvatures are assumed to be the same for each lamina; loading does not affect laminate thickness.
The differential composite element (Fig. 32.2) is described in terms of curvilinear coordinates a, p and y. The coordinate axes a and p coincide with lines of principal curvatures of the basic surface located at a distance e measured from an outer element surface. The geometry of this basic surface is defined by the coefficients of the first quadratic form A,, A , and the principal radii of curvature R, and R,. Therefore, the governing equations of composite structural analysis become as follows:
= 0 (1, 2 )
General equations for composite structures 739 where 1
dA,
(L2) A 1 4 aa The symbol ( I I 2) indicates, hereafter, that an equation written for the a variable yields one more analogous equation for the /3 variable by 1 and 2. commutating the indices a, /I, The foregoing set of equations (32.1, 32.2 and 32.3) include 23 equations and the same number of unknowns, namely: 0 the stress resultants N,Q and the couples M acting on the basic surface shown in Fig. 32.3; 0 displacements ua, up/ w of the points on the basic surface in the a,p and y directions; 0 rotations ea, 8, of the normal to the basic surface; 0 rotations of the tangents ma and osto the basic surface in the cry and ,by planes; 0 strains E and flexural deformations K of the basic surface; 0 transverse shear deformations 3, and $in J, the ay and ,by planes. a1 = _ _ _ _
and F = (N/ M). 2. Constitutive equations
’t 3. Strain-displacement equations & a
1 au, w 1 + a u +-+-w; A, aa P R, 2
(1,2)
=----
1 auB
A , aa
1
Kap
1
aua ap
+---ulu,-au
Eap = --
A,
ao,
= -~ +
A , aa
1
1 2 8
+-w 2 wa
P
ae, + u p a + u2ep
”P
--
A, $3
Fig. 32.3 External and internal forces and moments.
0
a
=
1 aw
R,
A, aa
(1,2)
(32.3)
The equations of motion (32.1) contain inertia terms with time derivatives of displacements and rotations; the equations reduce to static
740 Analysis methods equilibrium without these derivatives. The coefficients B and Dpcharacterize the inertia properties of the element in Fig. 32.2 with respect to displacements and rotations. C,, on the other hand, is the coupling coefficient describing the interaction between displacements and rotations.
where the Fs are the body forces and p , 9 are the simultaneously applied surface forces as shown in Fig. 32.3. Functions 7 and 6 reflect variation of the wall thickness, h (a,B) = e + s, along the coordinate lines, i.e.
and e (a,p) It is assumed that functions s (a,/I) vary rather slowly, therefore sin 7 = 7,cos 7 = 1, sin 6 = 5 and cos 6 = 1. It must be noted that, Fig. 32.4 Layer coordinates. for variable thickness, the layer coordinates ti in Fig. 32.4 can also depend on a and p. For For a thin walled laminate (Fig. 32.4) these structures with constant thickness, 7, = l;lp= 0 and Ea = 6,= 0. The governing set of equations coefficients become (32.1, 32.2 and 32.3) accounts for geometric B = I(0) P P nonlinearity and is applicable for small strains and moderate rotations o,(sin w = w ,cos w = 1) C = I(1)P P P of the structural element in Fig. 32.2. For thinwalled composite structures, the first-order D P = I (P2 ) - 2eI P(1) + $1P(0) (32.4) nonlinear effects, associated with variation of where (r = 1,2) radii R, and R2 under loading, can play an important role; in this case, the underlined rotations w are omissible when compared with their derivatives describing the changes in the p is the material density, and ti- ti are coordi- curvature of the basic surface under loading. nates of layers. The external loads in (32.1), The linearized buckling equations are formureduced to the basic surface, are specified by lated by modifying eqns (32.1), (32.2) and the following relations: (32.3) using the following assumptions: h 0 equation (32.1) excludes the terms corresponding to inertia, body and surface fa = Fadt + Pa + 9, + + 0,) + 9(7, - u p ) 0 forces; 0 in nonlinear terms of eqn (32.1), N, = N :, (112) N,= Np”, Nap= N u t where N,O, Np”, Nap”are the stress resultants that correspond to membrane prebuckling state of the structure; 0 constitutive equations (32.2) should exclude terms W and Mr.
1
ma
General equationsfor composite structures e Strain-displacement
equations exclude non-linear terms.
(32.3)
741
and (32.10) depend on the thermo-elastic constants of the material. For a composite layer-referred to principal material coordiThe membrane (B,,), the bending (D,,), the nates 1, 2, 3 in Fig. 32.5, with axis 1 enclosing coupling (Cmn),and the shear (Kmn),stiffness an angle 'P, or -ql with the a axis as in Fig. 32.6. coefficients in the constitutive equations (32.2) These coefficients are are determined with the following equations: A,:) = E ([IC 4 + EC')s4 + 2E ( 1 ) ~2s 2 1 1 2 [ 12 B mn = I mn (0) 1
Cmn = 1:)-
A12(1)= A 0) =
el::)
21
yl$)Ey)+
A$] =
(Et) + EO - 2El$))c;s;
+ E~]c:+ 2E 12( ' 1 ~ 2 ~ 2 1
f K = K4Ks5-K4;
(32.6)
where (mn = 11,12,13,23,22,33; r = 0, 1,2)
-
l _ r +_l
1k
(32.7)
AtA(t;+ )::;t
i=l
and (rnn = 44,45,55)
z h 1 Kmn= Knm = - a dt = h210 mn h2
k
Fig. 32.5 Element of a unidirectional ply.
il
a{Ahi
(32.8) . , Constitutive equations, (32.2), include also temperature terms
N,' = J:) M,' = J:) - e]:)
(32.9)
where (m = 1,2,3; r = 0 , l )
The stiffness and the temperature coefficients of the ith layer, A,,, Am,in eqns (32.7), (32.8)
I
Fig. 32.6 Angle-ply laminate.
1
thermal strains and stresses are taken to be zero. The constitutive equations (32.2) also allow for hygrothermal effects if AT) = &'1A€li and A$) = @)AHi where AHi is the change of moisture concentration in the ith layer of the material. For an orthotropic layer! with the orthotropy axes coinciding with the a and the /3 axes, the nonzero stiffness and thermal coefficients are as follows:
ci = cos pi
var
=
fi) pa
where Es (s = 1,2,3) is the modulus of elasticity in the sth direction; G,, is the shear modulus in the st-plane; v,, and v,,are Poisson's ratios satisfying the following symmetry condition Ey,, = E, vts;alTand aZTare thermal expansion coefficients; ATi is the temperature change of the ith layer measured from the value at which Fig. 32.7 A sandwich wall.
t
= "'i
General equationsfor composite structures 743 For a sandwich wall with a light core and thin facings as in Fig. 32.7, the transverse shear deformation of the facings can be negligible in comparison with the corresponding deforma= 03) . On the tion of the core (AZ’W = 03, other hand, the in-plane stiffness of the core can be ignored in comparison with the corresponding stiffnesses of the facings (A): = A 12 = A,?) =A (2) = 0). 33
b
Therefore, eqns (32.7) and (32.8) yield k
h, K mn = -a h2
(2)
mn
As a first approximation a system of densely arranged equivalent ribs can be also treated as a system of continuous layers as depicted in Fig. 32.8. For the layers, simulating the rib elements in Fig. 32.9, the non-zero inplane stiffness coefficients are A:) = E-d
Fig. 32.9 Rib parameters. ii
A, =
cjEjsin4pj j=1
n
A,, = A,, = A,
=
Ill
cjEisin2pjcos2pj 1=1
c.G.cos’pj
j=1
d
n
rn
A$) = E-C d
A$) = E-b d
where it is assumed that the ribs are parallel to the a axis, and E is the rib modulus. For a lattice structure in Fig. 32.10:
C n
A,, =
cj~ps4pj
j=1
Fig. 32.8 A stiffened wall.
A,, =
cjGjs’n2p, j=1
where cj = di/uj, Ej and G. are the elasticity and the shear moduli for the jth system of ribs. The stiffness coefficients in the constitutive equations (32.2) are associated with the following deformations of the composite structure: (a) Bll, B,,, B,, stretching and contraction of the basic surface due to corresponding loading or Poisson’s effect; shear of the basic surface; @) 4 3
744 Analysis methods
JP Fig. 32.10 Lattice composite structure.
(c) D,,, D,,, D,, bending of the basic surface These equations (note that mn - 11, 12, 22, 13, due to corresponding loading 23,33) can be satisfied for the following strucor Poisson’s effect; tures: (4 D33 twisting of the basic surface; homogenous or symmetric with respect to (e) K1,, K2, transverse shear of the wall in the middle surface of thin laminates for the a y and by planes; which e = s = h / 2 (see Fig. 32.2); (f) B13, B23 stretching-shearing coupling laminates consisting of isotropic layers with deformation (a) and (b); different moduli of elasticity (E,) and thick(g) Cll/ C,,, C,, bending-stretching coupling ness (hi)but the same Poisson’s ratio, vI= Y, deformation (a) and (c); for which (h) c3, shearing-twisting coupling k k n deformation (b) and (d); (i> ‘23 stretching-twisting (a) and (d) i=l i=l and shearing-bending (b) and sandwich structures with facings made (c) coupling deformations; from the same material but having different (j) D13,D, bending-twisting coupling thicknesses (see Fig. 32.7) for which deformation (c) and (d); 1 (k) K,,, KZ1 interaction of transverse shear e= [h,2+ h3(h3+ 2h, + 2h,)] deformations in the a y and by 2(h, + h3) planes. In the general case, eqns (32.12) are incomThe coordinates of the basic surface, e, in eqns patible, because it is impossible to find a basic (32.4), (32.6) and (32.9) can be taken arbitrarily surface for an arbitrary stacking sequence and as a rule, is used to eliminate coupling such that all the coupling stiffnesses simultastiffnesses Cmnin the constitutive equations. neously become zero. However, one of the coupling stiffnesses can always be eliminated Thus, in accordance with eqn (32.6): with the aid of the corresponding equation in Cmn = lmLl)- elmlo)= 0 (32.12) eqns (32.12).
Composite beams 745 There exists the so-called method of reduced bending stiffnesses according to which each of eqns (32.12)is used to obtain the corresponding coordinates - 1 (1)/I (0)
emn-
mn
mn
Then, eqns (32.6) yield
B mn = Imn(0)
eap
(a,PI Y)= Eap (a,P) + Fap(a'P)
Strains in the k p, plies, whose stiffnesses are specified by eqn (32.11),have the form
cmn =0
e?) = e$ cos2pf + ep(l) sin2p, The method of reduced bending stiffnesses is not formally established and can give both satisfactory results and large errors. Also, it can be noted that eqns (32.12) can sometimes be satisfied if there exists a possibility to change the stacking sequence ot the layers, i.e. to use coordinates ti as unknown factors. The governing set of equations, (32.1), (32.2), (32.3), is of the tenth order with respect to variables a and P; five boundary conditions need to be formulated at each point of the boundary. If the edge of a composite structural element is clamped, then
u = u =w=O a
P
a
P
rt
e$) sin p, cos p,
elit)= f (e:) - e:')) sin 2pl + e,:)
cos 2pf
where, e.g. ea(') = ea (a,P, y = yl). The stresses, in terms of the principal material coordinates of a ply (see Fig. 32.5), are as follows: 0:)
= Ey)(el(l)+ y 12(')e2( 1 ) ) (1, 2)
r12(i) = G (')e ( i ) 12 12
To evaluate the strength of the ply, these stresses should be substituted into any available strength criterion.
=o
For a simply supported edge a = constant (1,2)
N, = up= w = M a = Op = 0 (1,2) For a free edge a = constant (1,2)
N, = Nap= Sa = M a = M a , = 0 (1,2) where
Sa = Q, - Nama - Napup (h2) For a linear problem, S, = Q,. The solution of the governing set of equations, satisfying the proper boundary conditions, specifies generalized strains -E, K, 11) and displacements u,,up, w, 8 , 8, as functions of a and P. Then, the displacements of any point of the structure can be found in terms of a, B and y by the following expression:
32.3 COMPOSITE BEAMS
32.3.1 LINEAR BENDING AND AXIAL LOADING
Combined linear bending and tension (or compression)of a composite beam in Fig. 32.11 are described by the following equations: N' = 0
M' = Q Q'+jT= 0 N = Bu'
M = DO'
746
P Fig. 32.11 Composite beam.
bk
Q = K(8 + V')
(32.13)
where ( )' = d( )/dx, j7 = pb, - qb,. The axial and the bending stiffnesses of the beams are
B
= Io
D = I2 - el, where e = Il/Iois the coordinate of the neutral axis. For a laminated beam with a cross section shown in Fig. 32.12
Fig.32.12 Layer coordinates.
M = M,
( n = O , l , 2)
8 = 8,+-
M0x D
and the transverse shear stiffness is defined by
1 V = Vo + -(Q$ K
Successive integration of eqn (32.13) yields the following general solution:
-88-
N = No NO B
u = uo + -x
Q = Q O - Q, - QR
~
+ Q,x f-
-Mp -MR
Qox' 2D - 0 -P 0 , - MP- MR)
Mox2 Q0x3 + vp+ VR 20 60 ~
where quantities with subscript '0'correspond to the initial cross section at x = 0. The following integral terms: QP = Iljidx
Thin-walled beams 747 force, should be taken into account. The corresponding equations describing such a type of beam behavior have the following form:
M' = Q Q'
Vp = I j p d x
N = B[u'
account for the distributed loads (see Fig. 32.11) and these next terms (Qp MR,OR,
PQ,",M,",e,", v,")
VR> =
+ NV" + jj = 0 + %(V')2]
M = DO'
Q = K(8 + V')
describe the action of the concentrated forces. where N is a constant axial force. If the beam For x < xm, the forces should be Q," = M," = 8," has fixed ends, N is determined from the = V," = 0 and for x 2 xm boundary conditions. For combined axial loading and bending, N is a known applied Q," = E m load.
M," = Ern(. - xm) 32.3.3 BUCKLING
-
Rm 0,m = - ( x 20
- xm)2
-
Rm VRm = - ( x 6D
- xm)3
Nc =
Rm= xm- F~ Displacements and stresses at any point of the beam are defined by
q x , y > = 4 x 1 + ijw
[
N +M(x) ax(x,y) = Ex 7 D
qsy) =-
~
E: )
The critical magnitude of an axial compressive force, causing the column to buckle, can be determined as NE 1 + (N,/K)
where, NE= c7c2 D/L2 is the critical Euler force, L is the column length, and c is the coefficient dependent on boundary conditions. For a colu r n with simply supported ends c = 1, for a cantilever column c = 1/4, and for a column with clamped ends (c = 4). 32.4 THIN-WALLED BEAMS
loy€xbijdy
ij = y - e
32.3.2 NONLINEAR AND LONGITUDINAL BENDING
For nonlinear transverse or longitudinal bending, the axial strain, due to the large lateral deflection developed by the applied axial
Composite thin-walled beams are used as members of trusses, aircraft propellers, helicopter rotor blades, drive shafts, etc. For an orthotropic beam whose cross section with two axes of symmetry (Fig. 32.13), the normal stress resultant nZ (Fig. 32.12) due to combined axial loading and bending has the following form:
(2 X )
nz=B-+-y
748 Analysis methods
Fig. 32.13 Thin-walled beam.
X
f(
Fig. 32.14 Normal and shear stress resultants acting in a thin-walled beam.
where
S = $Bds
D = §By2ds
uy(z,s)= v(z) - xOz(z)
and B is the axial stiffness of the beam wall. For a beam with a non-deformable cross section contour B = Bl1. For a beam whose cross-section contour can be treated as absolutely compliant in its plane B = B,, B1,2/B,,, where Bmn are determined with eqn (32.6). The shear stress resultant nzs (Fig. 32.14) developed by a transverse force and a torque is
uZ(z,s)= w(z) + yex(z)+ f ( s )
(32.15)
where o, w are the displacements in the y and z directions and Ox, Oz are the angles of rotation about the x and z axes of the beam cross section. These displacements are defined as follows: '0
=
:I
v,, + (V,, - Bx)dz
where
]
D SX(s) =
I
(32.14)
By ds
where A is the area bounded by the contour of the cross section and Y is the length of the normal to the contour shown in Fig. 32.14. The displacements of an arbitrary point on the beam cross-section contour (Fig. 32.14) along the x, y and z axes can be expressed as u x W= Y q z )
ez = e; + c z k z d z where o,, w,,Ox0,O: are the displacements and the rotations of the initial cross section at z = 0, cz =
-9-1
ds
4A2 B33 and V,, is the beam shear deformation having the following f01-m:
Rectangular plates 749 F (s) is specified by eqn (32.6) and B,, is the d e a r stiffness of the beam wall given by the corresponding equation in (32.6). Function f(s), entering the last equation in (32.15), determines the free warping of the beam cross section under torsion and bending and has the form
32.5.1 LINEAR BENDING OF SYMMETRICALLY LAMINATED PLATES
If the laminate has a symmetric lay-up, the basic surface is at e = h / 2 and eqns (32.6) and (32.7) yield
Bmn = 2
c
A m n Y zz zZ-J
1=1
cmn= 0 Note that the ),‘( and the transverse (Q,) as as the bending moment (Mx) acting on the beam cross sections and entering the foregoing equations - are governed by the equilibrium equations for a beam element similar to equation (32.13). 32.5 RECTANGULAR PLATES
Composite plates possess high specific strength and stiffness. They are used as structural elements for aircraft, ships and other structures. The governing equations for a plate (Fig. 32.15) can b e obtainedin Cartesian Soordinates from eqns (32.1), (32.2), (32.3) if we take A, = A, = l , l / R , = 1/R, = 0 and replace a, PI Y,with x,y, z.
i=l
where z i are the layer coordinates shown in Fig. 32.16. Then, the problem of bending of an orthotropic plate under the combined action of surface pressures p and q (see Fig. 32.3) is reduced to the following equation in terms of plate deflection w (note that the transverse shear deformation is ignored): a4w
+ 2 ( q 2+ 2 4 D1lF axzayz
where
a4w
a4w
=
+
’22-
af
=F)
- 4. The strains at an arbitrary point of the plate are defined
kl2
i
i Y Fig. 32.15 Rectangular plate.
Fig. 32.16 Layer coordinates for a symmetric laminate.
750 Analysis methods
where
k, = Wi (X = a/2) they allow to find strains and stresses in terms k, = W. (X = ~ / 2 ) of principal material coordinates of the composite plies. k, = W;’ (X = a / 2 ) The variational approach, based on the approximate analytic expression for the i = 1, 2 and ()’ = d ()/dx. Functions W,(x) and deflection in the y-direction while satisfying W,(X) are particular solutions that are symmetthe corresponding boundary conditions, ric with respect to coordinate x and satisfy the yields the following approximate (but rather following ordinary differential equation: accurate) solutions for a plate (Fig. 32.15) w”” - 2s2w” + p w = 0 loaded with a uniform pressure jf = po = constant. The approximate deflection equations where for different boundary conditions are as folc2(D,, + 2033) lows: 52 = C P l l
(a) Plate with simply supported longitudinal and transverse edges W(XIY> =
wp [1 -&)lP,(y)
(32.16)
(b) Plate whose longitudinal edges (y = 0 and
y = b) are simply supported and transverse edges (x = * a/2) are clamped W(X1Y) =
wp 11-f,(X)lP,(y)
c parameters in these equations have the form b
c1 = p d Y
(32.17)
(c) Plate with clamped longitudinal and simply supported transverse edges
b
c2 = l0(PWY b
wp [1 -f,(41P2(y) (32.18) (d) Plate with clamped longitudinal and transverse edges W(XlY> =
w(x,y) = wp [1 -f2(X)IP2(y) (32.19) The following are used:
c3
= l0(Pf7ZdY b
c = ITdY For solutions in eqns (32.16) and (32.17) it should be taken c, = 0.04921 b9, c, = 0.48571 b7, c3 = 4.8b5, c = 0.2 b5. For solution in eqns (32.18) and (32.19) c, = 0.001587 b9, c2= 0.01905 b7, c3= 0.8 b5, c = 0.03333 b5. 32.5.2 BUCKLING AND POST-BUCKLING BEHAVIOR OF SYMMETRICALLY LAMINATED PLATES
In-plane compression or shear (Fig. 32.17) (where the forces Tx, T,, Tx, are uniformly
Rectangular plates 751 (b) if the edge y = 0 is simply supported and the edge y = b is free, then
Under pure shear (Tx= Ty = 0), the critical load for an orthotropic symmetrically laminated plate is expressed as 7e
Tx; = k--.\j(D,,D,) ab Fig. 32.17 In-plane loading of a rectangular plate.
(32.22)
Coefficient k is given in Table 32.1 for typical values of the following parameters:
distributed along the plate edges) can result in plate buckling. For a simply supported rectangular (a 2 b) plate under uniaxial compression (T, = Txy= 0), the critical load is expressed as Since the value of the critical load for an (32.20) orthotropic plate does not depend on the direction of shear forces, the parameter p in where Table 32.1 can be replaced by 1/ p , so that Table 32.1 presents coefficients k as p varies from k = 2 1 + D12 + 2033 (32.21) 0.02 to 50. Critical combination of compressive and Note that eqn (32.21) is valid if the longitudi- shear forces can be determined using the folnal edges of the plate (y = 0 and y = b in Fig. lowing equation: 32.17) can experience displacement along the 2 T X =I y-axis. If these edges are fixed in this direction (which is often the case), then the compressive forces Tx give rise to transverse compressive where T,' and Tx; are specified by eqns (32.20) forces Ty = vq T, due to Poisson's effect. Then and (32.22). Used as the skin elements of stringer panels [DllA + 2(D1, + 2D,,) +-k = or shear webs, composite plates (just as metal A 1 + (vx,/4 ones) can sustain high compressive or shearshould be substituted into eqn (32.20) and ing loads after buckling. However, in contrast minimized with respect to 1 = (rnl/a)2 where rn to metal panels whose ultimate loads are usuis the number of half waves in the x-direction. ally determined by rib fracture, buckling The following approximate expressions are failure of composite panels (particularly made derived for the k coefficients with different from carbon-epoxy composites) is often plate edge supports: caused by skin fracture due to bending. (a) if the edges y = 0 and y = b are clamped, Therefore, traditional engineering methods of strength analysis, such as the method of then reduced width for compressed panels and the Dl2 + 2033 concept of diagonal stress field for shear webs, can hardly be used for composite panels; more
T,' = k--./(D,,D,) 7c2 b2
) TvJ
q+[+]
752 Analysis methods Table 32.1 Buckling coefficient k for pure shear
P q 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.2 3.4 3.6
0.02 0.04 25.5 28.9 32.2 35.3 38.4 41.3 44.1 46.8 49.4 51.8 54.2 56.4 59.3 62.1 67.8 70.6 73.4
18.4 20.8 23.1 25.4 27.5 29.6 31.5 33.4 35.2 36.9 38.5 40.1 42.1 44.0 48.0 50.0 52.0
0.06 15.3 17.3 19.2 21.0 22.8 24.4 26.0 27.6 29.0 30.3 31.6 32.8 34.4 36.7 39.3 40.9 42.5
0.08 13.5 15.2 16.9 18.8 20.0 21.4 22.8 24.1 25.4 26.6 27.7 28.8 30.8 31.4 34.2 35.6 37.1
0.10 12.4 13.9 15.4 16.8 18.1 19.4 20.7 21.8 23.0 24.0 25.0 26.0 27.1 28.3 30.8 32.1 33.4
0.14 10.9 12.2 13.4 14.6 15.8 16.9 18.0 19.0 20.1 21.1 22.0 23.0 24.1 25.3 27.4 28.5 29.5
0.18
10.4 11.5 12.4 13.6 14.7 15.7 16.7 17.7 18.6 19.5 20.4 23.1 22.3 23.3 25.1 26.0 26.8
rigorous solutions of the corresponding nonlinear problems are required. 32.5.3 NONSYMMETRICALLY LAMINATED PLATES
In contrast to symmetrically laminated plates, bending of plates with an arbitrary stacking sequence of the layers is accompanied by stretching of the basic surface. Also, the plate deflections depend on boundary conditions imposed on the inplane displacements2. For in-plane compression, nonsymmetrically laminated plates experience bending which should, in general, be described by nonlinear equations. This longitudinal bending can be unstable and can be usually accompanied by the so-called mode-jumping. 32.5.4 AXISYMMETRIC DEFORMATION OF
CIRCULAR PLATES AND DISKS
The problem of axisymmetric bending and inplane deformation of orthotropic composite plates and disks (in cylindrical coordinates Y,
0.24
0.32
0.40
0.50
0.60
0.80
1.00
9.61 10.4 11.5 12.4 13.3 14.2 15.2 16.0 16.8 17.8 18.4 19.3 20.3 21.0 22.8 23.7 24.5
8.40 9.27 10.2 11.0 11.9 12.8 13.6 14.5 15.4 16.4 17.1 17.9 18.8 19.5 21.3 22.1 22.9
7.50 8.36 9.24 10.1 11.0 11.8 12.7 13.7 14.4 15.2 16.1 16.7 17.9 18.7 20.4 21.2 22.0
6.77 7.66 8.52 9.39 10.4 11.2 12.1 12.9 13.7 14.6 15.4 16.2 17.1 17.9 19.7 20.4 21.2
6.32 7.20 8.08 9.05 9.93 10.8 11.6 12.4 13.3 14.1 14.9 15.8 16.7 17.4 19.1 19.9 20.7
5.87 6.73 7.68 8.55 9.45 10.3 11.1 12.0 12.9 13.8 14.6 15.4 16.3 17.1 18.7 19.6 20.3
5.76 6.66 7.56 8.47 9.34 10.1 11.0 11.9 12.8 13.7 14.5 15.3 16.2 17.0 18.7 19.5 20.2
p, z) is reduced to the following set of
equations in terms of radial displacement of the basic surface ut, deflection w,and rotation of the normal to the basic surface 6 , i.e. ?ut’”’
+ ~ Y ~ U +~ (7 “ ’ - np2 - n2)Yur” + (1 - np2
- n:)ur’
-
I
+ (n?:
- n 4)2 =‘ [ -(C ‘ 1 C
Y
Bll Dll
p’rdr) - Y ( Y F ~ )-” ( Y F ~ )+’ n;F,
1
(32.23)
1 In r Kll Here, ()’ = d()/dr, j7 = p - q (see Fig. 32.3), Fr(r) is a radial body force (e.g. a centrifugal force for a spinning disk), C,, C, are constants of integration, and
w
= -(Cll
Cylindrical shells 753 nb2
=
32.6 CYLINDRICAL SHELLS
D22 ~
Dll
n: =
___ c2;
BllDIl Coordinates of the basic surface e = I $ ) / I z ) provide, in accordance with eqn (32.12), the zero value of the radial coupling stiffness (Cll = 0). Radial and circumferential strains at an arbitrary point of the plate are expressed as
er = u,’ 1 ep = -(ur r
’,
a j
h
+ zOrf + z0,)
The general solution for eqn (32.23) has the form
1c p + 6
ur =
Filament wound composite cylindrical shells are used as pressure vessels, reservoirs, pipes, aircraft and ship elements. The governing equations for a cylindrical shell (Fig. 32.18) can be obtained from eqns (32.1)-(32.3) if we take = = l / R i = or ‘ 2 = and rep1ace 7 with yTz.
Fig. 32.18 Cylindrical shell.
up
1=3
where up is a particular solution and si are the roots of the equation
s4- (nP
+ n,’)s2 + np2n:
- n:
= 0
Six constants of integration can be found from the corresponding boundary conditions according to which ur = w = Or = 0 for a clamped edge, w = Nr= M , = 0 for a simply supported edge, w = ur = Mr = 0 for a hinged edge fixed in the radial direction, and Nr= Mr = Q, = 0 for a free edge. To write the force boundary conditions in terms of displacements, the following expressions can be used: Ur
O
Nr= Bllu,’ + B 1 2 T + C122 r
32.6.1 AXISYMMETRIC DEFORMATION
One of the most important loading cases for cylindrical shells is the axisymmetric loading with pressures p , q and axial forces N (Fig. 32.18). In this case, the equations account for the first-order nonlinear effects of the axial forces on the curvature of deformed shell meridian. These equations have the following form:
M,’- Q, = 0 N Q,‘+ Nw” - 2 + j7 = 0 R N,= N W N, = B,,u‘ + B l2R
U
0
M, = C12& +DllBr’+ D 12 r
_f_
W N, = B,,u‘ + B 22 + C,,Ox’ R
M, = C
W
-
l2
R
+ D,,OX’
754 Andysis methods
Q,
= K,,(ex + w')
(32.24)
where ( )' = d( )/dx, j? = p - q. Stiffness coefficients B, C, D, K are specified by eqn (32.6) in which e = I (,)/I(0) 11
Boundary conditions should be written in terms of w and force Sx = Nw'.
11
The foregoing set of equations, (32.24), can be reduced to the ordinary differential equation
32.6.2 NONSYMMETRIC DEFORMATION
In the general case of loading, composite cylinw""- 2s2w"+ PW = kp (32.25) drical shells can be usually described rather adequately by the so-called semimembrane where theory that, in addition to membrane theory, takes into account the circumferential bending s2 = C,, (1 + C ) + RN + ___ moments. The model of a semimembraneshell 2RCD can be represented by a system of rings with inextensible axes that take only circumferenB tial bending moments and by a system of t4 = R2B,,CD absolutely flexible beams that connect the rings and sustain axial and shear forces only. The semimembrane theory assumptions lead to the following equations: ~
=
'11['
+
C,, + RN RCK,,
1
aNx -+ax
a%, = o ay
a%, %+ay ax
+ - Q, +q R
Y
=o
B = B,,B, - B,:
Constants of integration entering the solution of eqn (32.25) can be found from the corresponding boundary conditions according to which w = Ox = 0 for a clamped edge, w = M x = 0 for a simply supported edge, and M x = S , = 0, where Sx= Q, + Nw' for a free edge. Thin-walled composite pressure vessels can be described by nonlinear membrane equations using the assumption that D,, = C, = 0. These equations can be reduced to W" - k2W = kp
Q,
where E
=
K,VY
a u w =-+-=(I
' a Y R
Cylindrical shells 755 8 Y
= q + - -v - aw R ay
(32.26)
where j7 = p - 9, 9Y is the circumferential surface traction, and stiffnesses B, D are specified by eqn (32.6) in which e = I$)/I$). Decomposition into Fourier series, i.e.
= Nx; = 0 for a free edge. The following expressions can be used for the boundary conditions Un =
BllR
,,,
‘11 wnr------w,
~
A,2B3:
N,”= BllR wn - R4 E1 , ~
m
B33
n=l
Nx; =
...! Ny”(x),My”@),p,(x)lcos Any
Wnr
-__
mnz
m t F 3 3
BllR
5 1 ~
m,3
Wnlfr
-
~
An’s3
I,
Z~~ w n’
where
m
It must be noted that the semimembrane theory is not valid for the case of axisymmetricloading whereAn= n/R, allows reduction of eqn (32.26) (n = 0) and reduces to the membrane theory for to the following governing equation: n = 1. Proper combination of solutions of eqn (32.25) for n = 0 and of eqn (32.27) for n 2 1 w ~ ” ”- 2s’~~’’ + t4wn= kp (32’27) allows consideration a wide range of practical where problems for composite cylindrical shells.
..., Qy”(x),qy”(x)lsinAny
52
=
A;(n2 - l)D, 2B33
32.6.3 BUCKLING
Under axial compression by forces N (here, in contrast to Fig. 32.18, N are compressive forces), cylindrical shells can experience three modes of buckling: column-type, axisymmetric, and nonsymmetric buckling. The actual critical load is the smallest of the three values. For a column-type buckling
Nc =
n2rn2R2B 2L2B2,(1 +
n2rn2R2B L2B22B33
)
where B = Bl1B2, - B1: and rn depends on the character of end fixity. If the end cross sections n are fixed in such a way that they can freely Four constants of integration entering the rotate (hinged column), then rn = 1.If the ends solution of eqn (32.27) can be found from the are clamped, then rn = 2. corresponding boundary conditions according Axisymmetric mode of buckling is typical to which u, = vn= 0 for a fixed edge (it is essen- for thick and sandwich shells. The corretial that inextensibility condition E = 0 yields sponding critical load is specified by the w,, = -RAnvn,so wn = 0 for a fixed edge) and N,” following equation that allows for transverse
Fn = P, -
9;
756 Analysis methods shear and radius variation through the shell
where
17 I=[."." BllD,
+=)I
+{(l
BP22
where
=(y) 2
1 ;
For a homogenous shell
Stiffnesses B, C, D are specified by eqns (32.6) and (32.7) in which e = I$)/Iio); Am,,should be where changed for Am,, where Ex= Ex/(l - vX,,vy,). (i) - A (1) A (11 = A (i), 421 = A ~ ( ~ ) / S ~ - i 11 12 12 The critical value of the lateral external presand sure can be approximated by 1 si = 1+ - (ti + ti-,-2e) 2R
41
Transverse shear stiffness is
Note that the shell is assumed to be simply supported at x = 0 and x = L (see Fig. 32.18). If transverse shear deformation is not taken into account, then
The critical load, corresponding to the general mode of buckling of a thin simply supported orthotropic shell, can be found as
N, = DllA:,,
+ D, 2 II", + R2Ai($+%]
Here, B,, and D, are specified by eqn (32.6) in which e = I$)/I$) and parameter c depends on the boundary conditions. For a simply supported shell c = 1,for a hinged shell whose end cross sections cannot move in the axial direction c = 1.5, and for a shell with one end hinged and the other end free c = 0.6. Buckling pressure for an infinitely long shell is
9, =
3D2,
x3
Finally, note that the derivation for equations presented in this chapter can be found elsewhere3.
References 757 REFERENCES 1. Tarnopolskii, Yu.M. and Kincis, T.Ya. 1985. Static Test Methods for New Van Nostrand Reinhold.
2. Whitney, J.M. 1987. Structural Analysis of Laminated Anisotropic Plates. Lancaster, Pensylvania: Techomic Publishing co,,Inc, 3. Vasiliev, V.V. 1993. Mechanics of Composite Structures. Washington: Taylor & Francis.
DESIGN ALLOWABLES SUBSTANTIATION
33
Christy Kirchner Lapp
33.1 INTRODUCTION
Designing with composite materials requires knowledge of a significantly greater number of properties than for conventional isotropic metals. The selection of lamina and laminate allowables can be critical in the analysis of a composite structure. However, composite design allowables may not always be obtained from a single source of data. Several references must often be consulted to determine all the properties in the necessary directions, especially if several fibers or matrices are being considered in the design. This can be a timeconsuming effort, especially during the initial design phase. In addition, some organizations may not have easy access to all the necessary references required to collect the data. This chapter assembles lamina data from numerous sources so that the engineer may have a single reference point for initial design and analysis of composite structures. A broad range of fibers has been included since composites are increasingly being applied outside the aerospace community. Design allowables in this section are for both elastic and strength properties. Elastic properties are necessary for laminate design or the analysis of composite structures. These properties include elastic moduli and Poisson’s ratios. Strength properties are required to predict laminate strengths or perform a failure analysis of the structure.
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Strength properties for compression, tension and shear must be determined. Both elastic and strength properties can be influenced by numerous variables such as the fiber, matrix, fiber volume and processing method. The test method used to determine the allowable can also affect the property. In addition, some allowables cannot be readily tested, especially properties through-the-thickness of the composite. Often, complete characterization of a fiber/matrix system may not be available and the engineer must estimate or assume properties. It can be expensive and time-consuming to completely characterize a fiber/matrix system during the initial design phase, so it is important for the engineer to have allowables based on reliable data and to understand the limitations. This chapter concentrates on providing the material database and techniques for assembling the necessary composite allowables for preliminary design. Final design allowables may require additional testing. The allowables provided in this chapter should be considered preliminary design values. Allowables for both two dimensional (2-D) and three dimensional (3-D) properties are included. As a starting point, lamina allowables for commonly used fiber/matrix systems are described and listed. The references for these allowables are included. Several references may be listed for a single material. Methods for estimating properties when data is not available or testing cannot be readily performed are also defined. The effects of processing methods on allowables and methods
Lamina allowables 759 for adjusting lamina properties for these vari- 33.3 LAMINA ALLOWABLES ables is described. The intent is to provide a Lamina allowables may be used in a laminated baseline for design allowables which can then plate code to predict laminate elastic and be expanded with additional testing data as strength properties, or they may be used required or modified for a specific application. directly in a finite element analysis code. Table 33.1 defines the lamina properties and the basis for each property. Most of the properties 33.2 NOMENCLATURE FOR DESIGN are based on test data for the 2-D properties. ALLOWABLES The 3-D properties typically represent calcuThe nomenclature used to describe composite lated values based on the equations shown in lamina and laminate properties is not consis- Table 33.1. tent within the industry. For this chapter, the Properties for commonly used fiber and lamina properties parallel to the fiber are spec- epoxy resin systems are included and the conified with a 1 and the lamina properties struction is assumed to be continuous fiber transverse to the fiber are designated with a 2. reinforcement. A wide variety of reinforcing Figure 33.1 illustrates this nomenclature for fibers is included since composites may be composite lamina and laminate properties. used in variety of applications, each with very The properties in the 1-2 plane are often different design requirements. Glass, polyethreferred to as the in-plane lamina properties ylene, aramid and graphite fibers are covered. whereas the properties through-the-thickness Graphite fibers include standard, intermediate (out-of-plane)of the composite are designated and high modulus types. The fiber/resin syswith a 3. The 1, 2 and 3 directions can be tems have very different properties and the referred to as the longitudinal, transverse, and actual composite application will dictate the through-the-thickness properties respectively. selection. For example, glass fibers are less expensive and more impact resistant than Laminate property definition
X
2
amina property definition
Fig. 33.1 Nomenclature for lamina and laminate properties.
760 Design allowables substantiation Table 33.1 Definition of lamina properties and equations used to calculate material properties
Lamina material properties
Equation used to calculate material property
Definition
Elastic El
E2 E3
GI2 G23
G13 v12
Elastic modulus in the fiber direction Elastic modulus transverse to the fiber direction Elastic modulus through-the-thickness Shear modulus in the 1-2 plane
Property based on test data Property based on test data Transverse isotropy: E, = E, Property based on test data E3
Shear modulus in the 2-3 plane
G23
2(1
=
+ YB)
Shear modulus in the 1-3 plane Poisson's ratio in 1-3 plane
Transverse isotropy: G,, = G,, Property based on test data
Poisson's ratio in 1-3 plane
Transverse isotropy: v13= vI2
Tensile strength in the fiber direction Compressive strength in the fiber direction Tensile strength transverse to the fiber Compressive strength transverse to the fiber Tensile strength through-the-thickness Compressive strength through-the-thickness Shear strength in 1-2 plane (in-plane) Shear strength in the 1-3 plane (interlaminar) Shear strength in the 2-3 plane (interlaminar)
Property based on test data Property based on test data Property based on test data Property based on test data u3 = u2 a3= a, Property based on test data Property based on test data
'23
'12 t13
'23
graphite fibers, but have a higher density and lower modulus. Since the lamina properties are assumed to be used as preliminary design and analysis parameters, the effects of temperature, environment and fatigue are not considered. However these conditions must be considered during the final design development. 33.3.1 TWO-DIMENSIONAL (2-D) LAMINA PROPERTIES
The elastic lamina properties required for a composite 2-D analysis are typically E,, E,, G,,, and vI2. The specific properties may depend
'23
=
'13
upon the type of analyses being performed or the analysis code being used. If a strength or failure analysis is performed, then the following strength allowables may be required; q, o,,-q,-0,and T,,. Tables 33.2 and 33.3 include these elastic and strength properties, which were compiled from various sources, such as military standards, material supplier data and published literature. References are included for each property so that the engineer may directly consult a particular reference if further information is required. These references are listed in Tables 33.4 and 33.5. The lamina properties for glass, polyethylene and aramid fibers are listed in Table 33.2.
Lamina allowables 761 Polyethylene and aramid fibers are more commonly known by their trade names as Spectra and Kevlar. Kevlar 29 and 49 are included for aramid properties. The polyethylene fibers include Spectra 900 and 1000. E-glass and Sglass (trade names) are included for glass fibers. Table 33.3 lists lamina properties for standard, intermediate and high modulus graphite fibers. Data for non-USA produced fibers is not included. The type of material system is also indicated in Tables 33.2 and 33.3. Design allowables for a prepreg material system versus a wet filament wound system may result in different properties. It is important to consider the processing method when selecting the properties to be used in a design. One processing method may result in a lower moduli or strength than another. The effects of different processing methods on design allowables is discussed in further detail at the end of this chapter. The majority of the 2-D properties is based on test data, not micromechanics equations. The test data is typically based on 'thin specimens' (typically less than 2.54 mm (0.10 in) thick). Properties based on thin specimens may not represent those for thick composites (typically greater than 6.35 mm (0.25 in) thick). The exact definition of a thick composite is not consistent within the composites community. Thick composite structures may have properties lower than those of thin composites, so the engineer may need to perform additional testing for certain applications.
these properties are difficult to test and data is not always readily available. The 3-D lamina properties listed in Tables 33.2 and 33.3 are estimated based on 2-D properties. The following section describes the methods for calculating lamina properties in the 3-direction when data is not available. 33.3.3 ESTIMATING LAMINA PROPERTIES WHEN DATA IS NOT AVAILABLE
Lamina properties through-the-thickness (3direction) are often not readily available, although they may be required to perform an analysis. When data is not available, these properties can be determined by assuming that the lamina is transversely isotropic. For a transversely isotropic lamina, the properties in the 2- and 3-directions are assumed to be the same. Thus the following equations may be used to determine elastic properties in the 3direction':
E, = E,
G*, = G12 r L3 G23
= 2(1 + vz3) '13
= '12
The transverse Poisson's ratio, v23, can be determined from the following relationship2:
33.3.2 THREE-DIMENSIONAL (3-D) LAMINA PROPERTIES
Irutially the majority of analyses performed on composite structures were two dimensional. However with advancement of finite element analysis programs and as new applications for composites arise, more analyses are being performed for the 3-D case. Thus through-the -thickness lamina properties have become necessary in performing certain analyses. Often
where vf is the fiber Poisson's ratio, V , is the fiber volume fraction, vm is the matrix Poisson's ratio, and Em is the matrix elastic modulus. Therefore to calculate certain properties in the 3-direction, the engineer needs to know certain fiber and matrix properties, and these have also been listed in Tables 33.2 and 33.3. For initial design purposes, it is simplest
762 Design allowables substantiation
to assume that the fiber and matrix are isotropic, although certain fibers are considered to be anisotropic. The fiber or resin shear modulus can be calculated by: E G=----'-2(1 + v) The following equations may be used to estimate strength properties in the 3-direction. u3= u2
-a3= -u2 The equations listed above should be considered a starting point for estimating 3-D lamina properties when actual test data is not available. The values may need to be verified by testing as the design progresses. 33.4 LAMINATE ALLOWABLES
Laminate elastic properties and strengths can be determined by testing or by using a laminated plate code. During the initial design phase, laminate allowables are typically determined by using a laminated plate code. This is especially true if the composite lay-up deviates from a 'standard' lay-up, such as a quasi-isotropic laminate ([90, +45, 4 5 , OIJ. There is typically more test data available for a quasi-isotropic lay-up than any other lay-ups. Laminate testing is often performed after completion of the initial design, material selection and composite lay-up has been decided. Laminate testing would be performed to confirm predictions and processing effects. 33.5 EFFECTS OF PROCESSING VARIABILITIES ON DESIGN ALLOWABLES
The actual fabrication method used to build a composite structure can have an impact on the design allowables. Some processing methods can result in a higher fiber volume fraction and lower void content than others. For example, an autoclave cured part using prepreg tape will typically have a higher fiber volume
and lower void content than a wet filament wound part. This can affect properties such as the elastic modulus (E,) or tensile strength (uJ. It is important that the engineer understand the limitations of the selected processing method and adjust the design allowables accordingly. Processing parameters which can affect the lamina allowables are fiber volume fraction and void content. Design allowables should be modified if the process used to determine the lamina properties deviates from the intended process for the final composite part; this ensures that unrealistic properties are not being used to design the part. There are numerous fabrication methods applicable for composites. The methods which will be discussed are those that apply to continuous fiber reinforcement. These include filament winding, hand lay-up and resin transfer molding (RTh4). The method of cure can also affect the lamina properties. For example, one part can be filament wound with prepreg tow and another can be hand layed up with prepreg tape. These parts have different fabrication methods, but they may be cured in the same manner; in an autoclave with vacuum and pressure. It is very likely that these two parts would have similar properties and require no adjustment of lamina properties. However, if the filament wound part was wet wound and cured in an oven without vacuum or pressure, then the lamina properties would need to be adjusted if they were based on properties derived from testing using prepreg tape. A wet filament wound part typically has a lower fiber volume fraction and higher void content. Thus in determining if lamina properties need to be modified, the engineer must consider the complete method of processing, including the raw material and cure method, not just the automated or manual process which is being used to fabricate the part. In general, if a part is wet filament wound, it will possess a lower fiber volume fraction and a higher void content than a hand-layed up part using prepreg tape and cured in an autoclave. Wet filament wound parts cured in
References 763 an oven without vacuum and pressure typically have a fiber volume fraction between 0.55 and 0.60 with a void content between 1 and 5%. Parts fabricated from prepreg tow or tape, which are cured in an autoclave with vacuum and pressure will typically contain fiber volume fractions between 0.60 and 0.65. Parts which are fabricated using the RTM process will typically possess fiber volume fractions of approximately 0.50. Determining lamina properties for RTM parts is particularly difficult since the preform is usually woven and properties are not readily available. Also weaving in some conditions may slightly degrade the properties. The adjustment of lamina properties should focus on the 2-D elastic properties; E,, E,, G,,, v,, and the 2-D strength properties; u,, -ol, u,, -pz and rl,. The simplest method is to adjust the desired property by multiplying the value by the ratio of the fiber volume for the selected processing method to the fiber volume listed in Tables 33.2 and 33.3. For example, if one has properties based on prepreg tape which has been autoclaved cured and wants to adjust these properties for a wet filament wound part the following calculation would be used to adjust the longitudinal elastic modulus (El):
E , (wet filament wound)
V,(wet filament wound) Adjusting properties by the ratio of fiber volume fraction is applicable for modulus, tensile and compressive strength, but does not serve well for Poisson’s ratios which would require micromechanics. The Poisson’s ratio can be calculated based on the following equation,: Y1,
=
v,v,+ vm(l - V,)
where vf is the fiber Poisson’s ratio, Vf is the fiber volume fraction and vm is the matrix’s Poisson’s ratio. REFERENCES Whitney, J., Daniel, I. and Pipes, B., 1984.
Experimental Mechanics of Fiber Reinforced Composite Materials. Brookfield Center, Connecticut: The Society for Experimental Mechanics. Vinson, J. and Sierakowski, R. 1987. The Behavior
of Structures Composed of Composite Materials. Boston: Kluwer Academic Publishers.
764 Design allowables substantiation Table 33.2 Lamina properties for glass, aramid and polyethylene fibers in epoxy matrices
Properties
E-glass/ DER 332
S2-glass/ DER 332
S2-glass/ xP251 s
Kevlar 29/934
Material system Fiber type Supplier Resin type Supplier Fiber volume Composite density, g/cm3 (ib/in3)
Wet wound E-Glass Owens Epoxy Dow 60% 2.05 (0.074)
Wet wound S-Glass Owens EPOXY Dow 60% 1.98 (0.072)
Prepreg S2-Glass Owens EPOXY 3M 60% 1.98 (0.072)
Prepreg Aramid DuPont EPOXY Fiberite 58% 1.38 (0.050)
51 (7.5) 17 (2.5) 17 (2.5) 7 (0.98) 7 (0.95) 7 (0.98) 0.25 0.32 0.25
54 (7.9) 5 (0.7) 5 (0.7) 2 (0.24) 2 (0.24) 2 (0.24) 0.40 0.47 0.40
Lamina elastic properties E,, GPa (psi x lo6) E,, GPa (psi x lo6) E,, GPa (psi x lo6)
G,,,GPa (psi x lo6) G,,,GPa (psi x
lo6)
G,,,GPa (psi x IOh) '12 '23 VI i
48 (7.0) 12 (1.8) 12 (1.8) 6 (0.84) 5 (0.70) 6 (0.84) 0.19 0.26 0.19
54 (7.9) 16 (2.3) 16 (2.3) 7 (0.98) 6 (0.89) 7 (0.98) 0.25 0.32 0.25
Lamina strength properties 01
(72
ff3
212
Tension, MPa (psi x lo3) Compression, MPa (psi x lo3)
1613 (234) 462 467)
1779 (258) -641 493)
1069 (155) -272 439)
Tension, MPa (psi x lo3) Compression, MPa (psi x lo3)
39 (54 -103 -05)
58 (8.4) -186 -(27)
9 (1.3) -130 -09)
Tension, MPa (psi x lo3) Compression, MPa (psi x lo3)
39 (5.6) -103 415)
58 (8.4)
9 (1.3) -130 -09)
MPa (psi x
lo3)
(3.3)
28 (4.0)
-186
427) 75 (10.9)
37 (5.3)
Tables 765
Kevlar 49/934
Kevlar 49 /DER332
Kevlar 149/934
Spectra 9OO/EPON 826
Spectra 1000/EPON 826
Pprepreg Aramid DuPont EPOXY Fiberite 58% 1.38 (0.050)
Wet Wound Aramid DuPont EPOXY Texaco 60% 1.35 (0.049)
Prepreg Aramid DuPont EPOXY Fiberlite 58% 1.38 (0.050)
Wet Wound Polyethylene Allied EPOXY Shell 55% 1.12 (0.040)
Wet Wound Polyethylene Allied EPOXY Shell 55% 1.12 (0.040)
72 (10.5) 5 (0.7)
5 (0.7) 2 (0.24) 2 (0.24) 2 (0.24) 0.41 0.48 0.41
82 (11.9) 5 (0.7) 5 (0.7) 2 (0.26) 2 (0.27) 2 (0.26) 0.31 0.38 0.31
106 (15.4)
6 (0.9) 6 (0.9) 2 (0.24) 2 (0.32) 2 (0.24) 0.34 0.42 0.34
31 (4.5) 4 (0.5) 4 (0.5) 1 (0.21) 1 (0.21) 1 (0.21) 0.32 0.40 0.32
50 (7.3) 1 (0.1) 1
(0.1) 1 (0.10) 0 (0.05) 1 (0.10) 0.28 0.36 0.28
1151 (167) -281 441)
12 (1.7) -134 419) 43 (6.3)
24 (3.5)
49 (7.1)
24 (3.5)
17 (2.5)
Continued on next page
766 Design allowables substantiation Table 33.2 continued Lamina properties for glass, aramid and polyethylene fibers in epoxy matrices E-glass/ DER 332
S2-glassl DER 332
S2-glass/ XP251S
Kevlar 29/934
MPa (psi x lo3)
66 (9.5)
66 (9.5)
77 (11.1)
34 (5.0)
MPa (psi x
66 (9.5)
66 (9.5)
77 (11.1)
34 (5.0)
72 (10.50) 0.09 33 (4.8) 3103 (450) 2.60 (0.094)
87 (12.60) 0.18 37 (5.3) 3792 (550) 2.49 (0.090)
87 (12.60) 0.18 37 (5.3) 3792 (550) 2.49 (0.090)
83 (12.00) 0.44 29 (4.2) 3620 (525) 1.44 (0.052)
3.4 (0.49) 0.35 1.2 (0.18) 64 (9.3) 1.22 (0.044)
4.1 (0.60) 0.35 1.5 (0.22) 83 (12) 1.30 (0.047)
Properties 213
r23
lo3)
Constituent properties Fiber E,, GPa (psi x lo6)
vf G, GPa (psi x IO6) Tensile strength, MPa (psi x lo3) Density, g/cm3 (lb/in3) Resin E , GPa (psi x lo6) 213
G , GPa (psi x lo6) Tensile strength, MPa (psi x lo3) Density, g/cm3 (ib/in31
3.4 (0.49) 0.35 1.2 (0.18) 64 (9.3) 1.22 (0.044)
3.4 (0.49) 0.35 1.2 (0.18)
64 (9.3) 1.22 (0.044)
Tables 767
Kevlar 49/934
Kevlar 49 DER332
Kevlar 149/934
Spectra 9OO/EPON 826
Spectra 1OOO/EPON 826
50 (7.2)
50 (7.2)
38 (5.5)
23 (3.4)
23 (3.4)
50 (7.2)
50 (7.2)
38 (5.5)
23 (3.4)
23 (3.4)
124 (18.00) 0.45 43 (6.2) 3620 (525) 1.44 (0.052)
124 (18.00) 0.28 48 (7.0) 3620 (525) 1.44 (0.052)
172 (25.00) 0.33 65 (9.4) 3448 (500) 1.47 (0.053)
117 (17.00) 0.30 45 (6.6) 2586 (375) 0.97 (0.035)
172 (25.00) 0.22 70 (10.2) 2992 (434) 0.97 (0.035)
4.1 (0.60) 0.35 1.5 (0.22) 83 (12) 1.30 (0.047)
3.4 (0.49) 0.35 1.2 (0.18) 64 (9.3) 1.22 (0.044)
4.1 (0.60) 0.35 1.5 (0.22) 83 (12) 1.30 (0.047)
2.8 (0.40) 0.35 1.0 (0.15) 83 (12) 1.30 (0.047)
2.8 (0.40) 0.35 1.o (0.15) 83 (12) 1.30 (0.047)
768 Design allowables substantiation Table 33.3 Lamina properties for graphite fibers in epoxy matrices
Properties
AS4/3501-6
Material system Fiber type Supplier Resin type Supplier Fiber volume Composite density, g/cm3 (ib/in3)
Prepreg Graphite Hercules EPOXY Hercules 60% 1.58 (0.057)
lM6/3501-6 Prepreg Graphite Hercules EPOXY Hercules 60% 1.55 (0.056)
IM7/3501-6
lM8/3501-6
Prepreg Graphite Hercules EPOXY Hercules 60%
Prepreg Graphite Hercules
1.57 (0.057)
EPOXY
Hercules 60% 1.58 (0.057)
Lamina elastic properties E,, GPa (psi x IO6)
143 (20.7)
159 (23.0)
159 (23.0)
186 (27.0)
E,, GPa (psi x lo6)
10 (1.4)
10 (1.4)
10 (1.4)
10 (1.4)
E,, GPa (psi x lo6)
10 (1.4)
10 (1.4)
10 (1.4)
10 (1.4)
GI,, GPa (psi x
lo6)
6 (0.85)
5 (0.71)
5 (0.72)
6 (0.80)
G,,GPa (psi x lo6)
3 (0.41)
3 (0.41)
3 (0.41)
3 (0.41)
G,,,GPa (psi x lo6)
5 (0.68)
5 (0.68)
5 (0.68)
5 (0.68)
0.30 0.52 0.30
0.30 0.52 0.30
0.30 0.52 0.30
0.30 0.52 0.30
'12
'23 '13
Lamina strength properties 01
Tension, MPa (psi x lo3) Compression, MPa (psi x lo3)
2172 (315) -1558 -(226)
02
Tension, MPa (psi x lo3) Compression, MPa (psi x lo3) 03
Tension, MPa (psi x lo3) Compression, MPa (psi x lo3)
59 (8.5) -186 427)
2413 (350) -1655 -(240)
2620 (380) -1862 -(270)
2689 (390) -1931 -(280)
Tables 769
T300/5208
T40/1962
T50/1962
P55/1962
P75/1962
Prepreg Graphite Amoco EPOXY Fiberite
Prepreg Graphite Amoco EPOXY Amoco
Prepreg Graphite Amoco EPOXY Amoco
Prepreg Graphite Amoco EPOXY Amoco
Prepreg Graphite Amoco EPOXY Amoco
62%
62%
62%
62%
62%
1.60 (0.057)
1.60 (0.058)
1.72 (0.058)
1.72 (0.062)
(0.062)
141 (20.5)
172 (25.0)
241 (35.0)
241 (35.0)
338 (49.0)
9 (1.2)
10 (1.5)
7 (1.1)
8 (1.1)
7 (1.0)
9 (1.2)
10
7
(1.5)
(1.1)
8 (1.1)
7 (1.0)
6 (0.92)
7 (1.00)
6 (0.84)
5 (0.79)
6 (0.85)
3 (0.41)
4 (0.54)
3 (0.39)
3 (0.39)
3 (0.36)
5 (0.68)
7 (1.00)
6 (0.84)
5 (0.79)
6 (0.85)
0.30 0.52 0.30
0.33 0.40 0.33
0.28 0.35 0.28
0.34 0.41 0.34
0.30 0.37 0.30
1524 (221) -1482 -(215)
3241 (470) -1724 -( 250)
1413 (205) -965 -(140)
931 (135) -510 474)
965 (140) 441 464)
36 (5.2) -159 -P3)
69 (10.0) -159 423)
37 (5.3) -159 423)
33 (4.8) -159 423)
33 (4.8) -159 423)
36 (5.2) -159 423)
69 (10.0) -159 423)
37 (5.3) -159 423)
33 (4.8) -159 423)
33 (4.8) -159 423)
Continued on next page
770 Design allowables subsfantiation Table 33.3 continued Lamina properties for graphite fibers in epoxy matrices
Properties 212
13
‘23
AS4/3501-6
lM6/3501-6
MPa (psi x lo3)
87 (12.6)
85 (12.3)
96 (13.9)
80 (11.6)
MPa (psi x lo3)
124 (18.0)
121 (17.5)
121 (17.5)
131 (19.0)
MPa (psi x IO3)
94 (13.6)
94 (13.6)
94 (13.6)
94 (13.6)
234 (34.0)
276 (40.0)
276 (40.0)
303 (44.0)
0.26 93 (13.5)
0.26 109 (15.9)
0.26 109 (15.9)
0.26 120 (17.5)
IM7/3501-6
lM8/3501-6
Constituent properties Fiber E , GPa (psi x lo6) Vf
G,, GPa (psi x
lo6)
Tensile strength, MPa (psi x lo3)
3930 (570)
5102 (740)
5309 (770)
5447 (790)
1.80 (0.065)
1.74 (0.063)
1.77 (0.064)
1.80 (0.065)
vr
4.4 (0.64) 0.36
4.4 (0.64) 0.36
4.4 (0.64) 0.36
4.4 (0.64) 0.36
G , GPa (psi x lo6)
1.6 (0.24)
1.6 (0.24)
1.6 (0.24)
1.6 (0.24)
1.26 (0.046)
1.26 (0.046)
1.26 (0.046)
1.26 (0.046)
Density, g/cm3 (ib/ i n 3 ) Resin Er, GPa (psi x IO6)
Tensile strength, MPa (psi x lo3) Density, g/cm3 (lb/ in3)
Tables 771
~
T300/5208
T40/1962
T50/1962
P55/1962
P75/1962
77 (11.2)
97 (14.0)
63 (9.2)
115 (16.7)
97 (14.0)
69 (10.0)
94 (13.6)
97 (14.0)
69 (10.0)
231 (33.5)
283 (41.0)
393 (57.0)
379 (55.0)
517 (75.0)
0.27 91 (13.2)
0.32 107 (15.6)
0.24 159 (23.0)
0.33 142 (20.6)
0.27 204 (29.5)
3241 (470)
5654 (820)
2413 (350)
1724 (250)
2069 (300)
1.77 (0.064)
1.80 (0.065)
1.80 (0.065)
1.99 (0.072)
1.99 (0.072)
3.9 (0.56) 0.35
3.7 (0.54) 0.35
3.7 (0.54) 0.35
3.7 (0.54) 0.35
3.7 (0.54) 0.35
1.4 (0.21)
1.4 (0.20)
1.4 (0.20)
1.4 (0.20)
1.4 (0.20)
66 (9.6)
66 (9.6)
66 (9.6)
1.27 (0.046)
1.27 (0.046)
1.27 (0.046)
50 (7.3) 1.27 (0.046)
66 (79.6) 1.27 (0.046)
772 Design allowables substantiation LAMINA PROPERTY REFERENCES
A
Weight density was calculated based on the following relationship:
D
Y~~was calculated based on the following equa-
tion:
pf = fiber density
Vf = fiber volume fraction p, = resin density
AAAS4 and IM6 fiber properties are based on DD Hercules data for IM7/3501-6 from the Graphite Fiber Products Handbook, based on mechanical Hexcel supplied data which was based on tow test data. test data for the fiber modulus and strength. The fiber Poisson's ratio was 'back' calculated E Assumed: based on the composite Poisson's ratio ( Y ~ , ) ,the o3= o2and *3 = +T~ matrix Poisson's ratio and the fiber volume EE IM7 and IM8 fiber properties are based on (see reference H). B L.L. Clements and R.L. Moore, Composite propHercules supplied data (Graphite Fiber Products Handbook) which was based on tow test data erties for E-glass fibres in a room temperature for the fiber modulus and strength. The fiber curable epoxy matrix, Composites, 1978, 9(2), 93-99. Properties for tI3were set equal to valPoisson's ratio was 'back' calculated based on ues for S2-glass/DER 332 since no data was the composite Poisson's ratio (Y~,), the matrix reported. Poisson's ratio and the fiber volume (see referBB Hercules supplied data for IM6/3501-6 from the ence H). Graphite Fiber Products Handbook based on F The following was assumed: mechanical test data. '23 = 31' C Composite is assumed to be transversely isotropic. A transversely isotropic composite is FF Hercules data for IM8/3501-6 from the Graphite a material which exhibits a special case of Fiber Products Handbook, based on mechanical orthotropy, whereby the properties are identitest data., cal in two orthotropic dimensions, but not the G Owens Corning fiber data and short beam third. The properties are the same in both shear data. transverse directions, but not in the longitudi- GG Properties based on test data listed in BASF nal direction. The following equations apply Hexcel Technical Information handbook. for transversely isotropic materials: H Equation for calculating the fiber Poisson's ratio (vf): E, = E, Y12- Y,(1
Vf
GI3 = GI2
1 '3
= '12
= 2(1
+
- VJ
Vf
E3 G23
=
Y2J
CC Properties were set equal to those for AS4 / 3501-6.
HH Amoco data for T300 fibers. Actual test method for fiber modulus and strength was not defined. The fiber Poisson's ratio was 'back calculated based on the composite Poisson's ratio using the equation in reference H. I Fiber and resin are assumed to be isotropic, therefore G, or GI is calculated as follows:
E* Gf = 2(1+
VJ
Lamina property references 773
R
I1
Properties based on test data supplied by Amoco for the T40/1962 system. Amoco supplied data for the 1962 resin system. Resin properties from publication by Texaco, J Huntsman Chemical Co. ’Jeffamine’, resin properties are based on 100 parts of epoxy resin and 45 parts of Jeffamine T-403. The DER 332 epoxy resin was cured with Jeffamine T-403. JJ Properties were set equal to T300/5208 value. KJK Amoco data for T40 and T50 from technical information sheets. Test method for fiber modulus and strength was not specified. The fiber Poisson’s ratio was ‘back‘ calculated based on the composite Poisson’s ratio (vlJ , the resin Poisson’s ratio and the fiber volume (see reference H). L H. Hahn, D. Hwaug, H. Chang, S. Lo, Flywheel Materials Technology: Design Data Manual for Composite Materials, UCRL-15365 Volume 1, P.O. 6641009, Lawrence Livermore Laboratory, July, 1981. LL Amoco test data for T50/1962 M Assumed
S
T
U V
u3= u2and -u3 = +2 MMAmoco test data for P55/1962. N Mil Handbook 17. NN Amoco data for P55 from technical information sheets. Test method for fiber modulus and strength was not specified. The fiber Poisson’s ratio was ‘back’ calculated based on the composite Poisson’s ratio (vlJ , the resin Poisson’s ratio and the fiber volume (see reference H). 0 Properties were set equal to S2-glass/DER332. 00 Amoco data for P75/1962. P Properties for XP251S epoxy were set equal to DER 332 since no data was available. PP Amoco data for P75 which was based on tow test data for fiber modulus and strength. The fiber Poisson’s ratio was ’back’ calculated based on the composite Poisson’s ratio (vI2), the resin Poisson’s ratio and the fiber volume (see equation D). Q DuPont supplied data for Kevlar 29/934, Kevlar 49/934, and Kevlar 149/934. Laminates were fabricated and tested by Boeing
W
X Y
Z
Technology Services, Boeing Commercial Airplane Co. No data was listed in DuPont literature for G12, therefore values for Kevlar 29/934, Kevlar 49/934, and Kevlar 149/934 were set equal to those for Kevlar 49. Hexcel rubber-toughened epoxy system based on a paper by S.R. Swanson, G.R. Toombes, and S.W. Beckwith, In-Plane Shear Properties of Composites Using Torsion Tests of Thin-Wall Tubes, 29th National SAMPE Symposium, April 3-5,1984. DuPont supplied data for Kevlar 29, Kevlar 49, and Kevlar 149 based on tow tests (ASTM D2343).The fiber Poisson’s ratio was ’back‘ calculated based on the composite Poisson‘s ratio (vIz), the resin Poisson’s ratio and the fiber volume (see reference D). Fiberite data for 934 resin system from Fiberite Material Handbook. Values for Kevlar 49/DER 332 set equal to those for Kevlar 49/Epoxy XD7575.03-XD7114Tonox 60-40 from Reference 6 of this list. D.F. Adams, R.S. Zimmerman and H.W. Chang, Properties of Polymer-Matrix Composites Incorporating Allied A-900 Polyethylene Fiber, SAMPE J., September/ October, 1985, pp. 44-48. Note: The modulus of Spectra composites is much lower than expected from the rule-of-mixtures relationship. A possible explanation is that the Spectra fiber modulus is a function of strain rate. For example a single fiber tested at 100%/min strain rate exhibited a modulus of 17 msi versus 11msi for a 8’%/min strain rate. H.W. Chang, L.C. Lin, A. Bhatnagar, Properties and applications of composites made of polyethylene fibers, 31st Intern. SAMPE Symp., April 7-10, 1986. t I 3for Spectra 1000 set equal to the value for Spectra 900. Hercules supplied test data for AS4/3501-6 determined by independent firms; Delsen Labs and McDonnell Aircraft Company. Hercules supplied data for 3501-6 resin. All data is listed in the Hercules Graphite Fiber Products Handbook. R.Y. Kim, E Abrams and M. Knight, Mechanical characterization of a thick composite laminate, Proc. Amer. SOC.Composites, 3rd Technical Conference, 1988, pp. 711-718.
774 Design allowables substantiation Table 33.4 References for lamina properties of glass, aramid and polyethylene fibers in epoxy matrices
Properties
E-glass DER 332
S2-glass DER 332
S2-glass XP251 S
Kevlar 29/934
A
A
A
Q
L L C
N N C 0
Material system Fiber type Resin type Fiber volume Composite density Lamina elastic properties
C C 0 D C
G23 G13 '12
'23 1'
?
Lamina strength properties (psi) ff1
02
Tension Compression
B B
L L
N N
Q Q
Tension Compresson
B B
L L
N N
Q Q
Tension Compression
E E
M M
M M
M M
B B F
L G F
N N F
Q Q
G H I G G
G H I G G
G H I G G
S H I S
J
J
J
J
P K I P K
T K I T T
0 3
212 '13
'23
F
Constituent properties Fiber E, f' Gf
Tensile strength Density Resin Er r'
GI Tensile strength Density
K I K
K
r
K
S
Tables 775
Kevlar 49/934
Kevlar DER332
Kevlar 149/934
Spectra 900/€ PO N 82 6
Q
L
Q
A
Spectra 1OOO/EPON 826
A
W W C
W C C W
D C
Q Q
L U
Q Q
V V
W W
Q Q
L U
Q Q
V
K
W K
M M
M M
M M
M M
M M
Q Q
Q Q
F
L L F
V V F
W X F
S H I
S
S
H I
H
V H
W H I W W
F
S S
S
I S
I V
S
S
V
T K I T T
J
T K I T
K K I K
T
K
K I
J K
K K I K K
776 Design allowables substantiation Table 33.5 References for lamina properties of graphite fibers in epoxy matrices
Properties
ASA /35016
lM6/ 3501-6
lM7/35016
lM8/ 3501-6
A
A
A
A
BB BB C BB
DD
FF
C DD
C
cc cc cc cc cc
C FF
Material system Fiber type Resin type Fiber volume Composite density Lamina elastic properties El E2
E3
G12
cc cc cc cc
G23
G,, v12 v2.3 31'
cc
cc
cc cc cc cc cc
Lamina strength properties (psi) 0,
Tension Compression
Y Y
BB BB
DD DD
FF FF
Tension Compression
Y Z
BB
cc cc
cc cc
Tension Compression
Z Z
E E FF FF
02
03
21' 31' 32'
cc cc cc
Y Y Z
BB BB
cc
E E DD DD
AA H I AA AA
AA H I AA AA
EE H I EE EE
EE H I EE EE
Y Y I Y Y
Y Y I Y Y
Y Y I Y Y
Y Y I Y Y
cc
cc
Constituent properties
Fiber Ef Yf
G*
Tensile strength Density Resin Er vr
Gr
Tensile strength Density
Tables 777
T300 5208
T40/ 1962
T50/ 1962
A
A
A
A
A
GG GG
I1 I1
LL LL
MM MM
00 00
C
C
C
C
L
cc cc
I1
LL
MM
C C
C C
C
L
cc
I1 D
LL D
C h4M D
C 00 C C
C
C
C
C
C
GG GG
I1 I1
LL LL
Mh4 MM
00 00
L L
I1
LL
MM
00
JJ
JJ
JJ
JJ
E E LL LL F
E E MM Mh4 F
E E
E E GG GG
cc
P55/ 1962
P75/1962
00
D
00 00
F
HH H I HH HH
KK H I KK KK
KK H
NN H
PP H
I KK KK
I NN
NN
I PP PP
GG K I GG GG
I1 K I I1 I1
11 K I I1 I1
I1 K I I1 11
I1 K I1 I1
II
MECHANICAL TESTS
34
Yu.M. Tarnopol’skii and VI L. Kulakov
34.1 STRUCTURAL HIERARCHY OF FIBROUS COMPOSITES
Fibrous composites are inhomogeneous materials with multiple levels of structural scale. The three levels of structural scale can be arranged in a hierarchy. The characteristic dimensions for the three levels are: fiber diameter, lamina thickness and plate thickness. The most appropriate test methods and structural analysis techniques are different for each level in the hierarchy. Test objectives and associated problems are also different for each level. The smallest scale is the diameter of the reinforcing fiber. The properties of the reinforcing fiber and polymer matrix and their interaction are studied in the field of micromechanics. The second level scale is the thickness of the unidirectional lamina. Macromechanics describes the properties of a monolayer under loading at an angle to the fiber direction. A monolayer is defined as a flat or curved element of material composed of a polymeric matrix and reinforcement of the same type and orientation throughout the layer. It is the basic structural element of laminated and fibrous composites. The characterization of monolayers by mechanical test methods is given particular emphasis in this chapter since testing of anisotropic materials is a relatively novel and seldom studied field of mechanics.
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
The test results are used to calculate the properties of materials with more complex configurations of fiber arrangement and of hybrids, i.e. materials with different reinforcements in the same lay-up. For multilayered composites, the largest structural scale is the thickness of the laminated plate that is equal to the sum of stacked laminae and interleaves. The theory of laminated plates allows one to determine the properties of the plate using the properties of the monolayers and their stacking sequence. If the monolayers are part of a structural element, then the highest level of scale is the characteristic size of the object. The dimensions of structural elements typically exceed the thickness of the plate by several times. The properties of components are measured by traditional mechanical and physical test methods’. The results of the analyses or tests at the first level are used as input for the analyses at the second level. The same relationship holds for the second and third levels. Upon converting each scale to a continuum at the next higher scale, it is important that for each level under consideration the number of elements be sufficient, e.g. through the width and thickness of the lamina or through the thickness of the plate, so the transition from a discrete medium to a continuum is performed without great error. This progression up the hierarchy eventually leads to solutions of real life problems involving structural elements or prototypes.
Problems of composite testing 779 34.2 PROBLEMS OF COMPOSITE TESTING
systems of coordinates are introduced: the axes of elastic symmetry in the material (1, 2, 34.2.1 HISTORY 3) and axes of loading (x, y, z for flat specimens; 8, z and Y for ring and tubular The first reference to mechanical testing of a specimens). It is preferable to use methods in structural material for an engineering applicawhich the x, y, z axes (or 8, z, Y ) coincide with tion is dated July 4,1662. The objective of the 1 , 2 , 3 axes. the tests was to compare the tensile strengths of The majority of laminated and fibrous comcords made of Riga and Dutch yarns. The posites exhibit low interlaminar shear and stronger in this contest was the thinner cord transverse tension strengths. Shear strength is manufactured in Riga. In the years that folcharacterized by the relations between E x / G x Z lowed, equipment and methods for testing (shear stiffness) and a;/zx; (shear strength). engineering materials, particularly metals, Transverse tension and compression strengths have reached a high degree of perfection and perpendicular to the fibers are determined by consistency. The appearance of composites E x / E Z , ax’u/az’u, axt”/a;, where Ex the relations and their ever-expanding use has once again and EZ are the moduli of elasticity in the x and made it necessary to improve mechanical test z directions; GxZ is interlaminar shear modulus; methods. Although significant progress has a,“ and a; are strengths in the x and z direcbeen made, there are vast differences in their tions; zxzuis shear strength in xz plane. The x maturity. The methods differ primarily in the and y axes are located in the fiber lay-up (reindegree to which they minimize extraneous forcement) plane, the z-axis is perpendicular stresses and strains. Although test methods to this plane; the (t) and (c) designate tension tend to become more complex as their accuand compression, respectively. racy increases, economics must be considered in their selection. Factors such as complexity of specimen preparation, amount of material 34.2.3 UNIQUE REQUIREMENTS OF required, and the requirements for specially COMPOSITE TESTING designed equipment must be considered. The anisotropy and unique structural properties of composite materials cause serious difficulties. For example, a large number of 34.2.2 DETERMINABLE VALUES strength and elastic properties must be deterThe purpose of mechanical tests is to deter- mined for complete characterization of the mine the strength and elastic properties of a material. Since the number of determinable material. However, only loads, displacements characteristics depends on the state of stress and strains can be measured in a mechanical and the degree of anisotropy23, one should test. The theory of elasticity for an anisotropic select the loading methods for which the body is used to determine the desired proper- experimentally determinable values are most ties of composites from these measurable simply related to the material characteristics. quantities. It should be remembered that The selection of techniques for analyzing the advanced fibrous composites with unidirec- data is critical as well as the determination of tional, laminated or spatial fiber lay-ups are their range of validity. Since the composite inhomogeneous, essentially anisotropic mate- analysis techniques are based on the theory of rials. The customary terms, i.e. tension, elasticity for an anisotropic body, it is necescompression, shear and bending, are meaning- sary to consider the error in treating an less without specification of the direction of inhomogeneous anisotropic medium as a conthe load and its relationship to the axes of elas- tinuous anisotropic medium. For example, the tic symmetry of the material. Therefore, two number of structural elements (fibers, lamina,
780 Mechanical tests etc.) must be sufficiently large to support this approximation43r6. Once the general test method has been selected, the details of the loading and the sample geometry must be selected. For fibrous composites, the principal difficulties lie in the generation of a uniform stress field in a representative volume of material, i.e. the elimination of end and edge effects. This is difficult even for the most simple types of tests. The difficulties increase with increasing degree of anisotropy, i.e. materials reinforced by high-modulus or high-strength fibers (boron, carbon and organic fibers). End effects are primarily influenced by the method of fastening and loading of the specimen, the length of the grip section, and the fiber orientation. The region involved in end effects extends in the direction of the greatest stiffness of the material and increases with the anisotropy of the material. Edge effects are primarily influenced by the size and shape of the specimen, the fiber orientation and the angle of specimen cutting. If strength anisotropy is present, improper loading and fastening can lead to changes in the failure mode and the resulting strength value. A most important considerationis the selection of the specimen width. The width must be large enough to avoid the effect of cut fibers at free edges which is important for specimens of off -angle, angle-ply and cross-ply materials. Edge effects are manifested as interlaminar stresses at the free edges of the specimen, the direction and magnitudes of which depend on the fiber lay-up. Material quality also causes unique requirements for testing composites. Quality cannot be ignored during testing because the material and structure are formed simultaneously. In addition, composites are extremely sensitive to mechanical and thermal history. Structural imperfections, in particular porosity, waviness and misalignment of fibers, require special attention. The presence of porosity affects the measurement of polymer matrix dominated properties, e.g. shear strength. Small amounts
of fiber waviness can cause the measured values of longitudinal modulus of elasticity and strength to be considerably lower than those of materials with ideally straight fibers. Fiber waviness also influences on the coefficient of thermal expansion in the fiber direction. The modulus of elasticity perpendicular to the fiber direction and the in-plane shear modulus are not significantly affected by fiber waviness. All of the aforementioned unique testing requirements apply to composites of a fibrous and laminated structure. Additional difficulties arise when spatially reinforced composites are tested because the transverse strength and stiffness are derived from a rigid framework rather than from a compliant matrix. 34.2.4 SUMMARY TABLES
The most common methods of testing composites in tension, compression, torsion and bending are described in Tables 34.1-34.5 . The high performance test fixtures designed specifically for composite testing, their description and recommended applications are given in Reference 7. 34.3 TEST SPECIMENS
The important relationships between fabrication methods, test methods and required specimen shapes are shown in Fig. 34.1. Specimens for mechanical testing are classified as flat specimens (bars and plates), rings (complete and segments) and tube@. The specimens and test methods in Fig. 34.1 are used to characterize the monolayer. Flat monolayers can be characterized with specimens that have a different fiber lay-ups but the same general, flat, long, narrow shape. To adequately characterize wound monolayers, it is necessary to use both rings and tubes. Ring specimens of a unidirectional fiber lay-up are used to assess characteristics in the fiber direction. Tubular specimens with a 90" wind angle are used to measure properties perpendicular
x
Ir
2h" .. I1
Test specimens 781
782 Mechanical tests
W
h
w
h
9
t l
J . 4
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%
Test specimens
784 Mechanical tests
m U
. Y3
m h aJ .r(
U
Y
2
d
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6
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a-
2
d
8
aJ
8
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0
m
&
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T
X
.r(
N
0
3
Y
.3
v
%
t
cn
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m u c\
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Test specimens 785
bn
@
786 Mechanical tests
f
t Fig. 34.1 Methods of material fabrication and respective specimens for testing.
methods, loading types, and failure modes are all different for the two test methods (Table 34.1). A uniaxial tension or compression test specimen has several functionally different parts: two loading sections, a gage section, and two transition sections. The loading sections provide a means of fastening the specimen in the testing machine. They receive and transmit the external loads to the gage section of the specimen. In the gage section, deformations are measured and stresses are calculated according to the geometrical dimensions and external load. The transition 34.4 TENSION AND COMPRESSION sections serve to attenuate stress-strain perturbations in the loading section to isolate 34.4.1 TENSION TESTING OF FLAT SPECIMENS them from the gage section. The specimen In spite of any analogy that may be drawn dimensions that are recommended in the between loading of flat composite specimens existing standards meet these requirements. in tension and compression, only the calcula- The specimen dimensions (length, width and tion relationships (taking into account the sign thickness) specified in standards as a function of the stresses and strains) are common to both of the type of fiber lay-up are shown in Table cases. The specimen shape and size, gripping 34.1 (Methods (a) and (b)). to the fiber direction. Tubular specimens with different balanced fiber lay-ups (fiber angles are symmetric with respect to the longitudinal specimen axis) are employed to assess shear characteristics and to study complex states of stress. The specimen shape, to a great extent, depends on the objective of the test: verification of scientific hypotheses, engineering specification of the material, or quality control of the materials. The most rigorous requirements are imposed on specimens of the second group.
Tension and compression 787 The greatest technical challenge in tension testing of composites, especially unidirectional composites, is the reliable transmission of tensile forces from the grips to the specimen. This is generally performed through the use of friction forces. Tabs bonded to the specimen improve the efficiency of load transmission considerably. The tabs should be made of a material that has a much lower modulus of elasticity and a higher total elongation than the respective characteristics of the specimen material. Tabs have been made of fiberglass reinforced composites, aluminum and wood veneers. The thickness of tabs should be between 1.5t and 4t, where t is the specimen thickness. The tabs must have a large enough area that the ultimate shear load capacity of the bond between the tabs and the specimen is greater than the breaking load of the specimen gage section. The mode of failure in tension depends on the relationship between the external load and the reinforcing fibers and on the type of reinforcement lay-up. When unidirectional composites are loaded in the reinforcement direction, they fail by breakage of the reinforcing fibers. This is accompanied by transverse cracks, longitudinal shear cracks and delamination of the polymer matrix. Increasing the angle between the load and the reinforcing fibers causes the mode of failure to change gradually from shear and splitting of the polymer matrix parallel to the fiber direction to pure transverse tensile cracking of the polymer matrix. The failure mode of composites with a balanced angle-ply reinforcement depends on the angle of the fiber lay-up. 34.4.2 COMPRESSION TESTING OF FLAT SPECIMENS
The main problem in compression testing of flat specimens is the selection of a loading method that ensures compressive failure. When the loading is achieved by normal
forces acting on the specimen ends (Table 34.1, Method (c)),it is impossible to achieve a sufficiently uniform stress distribution over the faces of the specimen. As a result, premature local failure of the specimen occurs. When the loading is achieved by shear forces acting on the sides of the specimen as specified by ASTM D3410 (Method (d)), the stress distribution in the specimen is also non-uniform, especially when flat wedge grips are used. The best method is a combination of the two methods in which normal forces are applied at the ends and shear forces are applied along the faces of the specimen grip section (Method (e)). In compression testing of unidirectional composites in the fiber direction, three basic modes of failure are observed: buckling of the reinforcing fibers, transverse cracking of the matrix, and shearing of reinforcing fibers at a 45" angle without local buckling of the reinforcement. Transverse cracking is caused by differences in the Poisson's ratios of the material components and by a non-uniform transverse strain distribution along the specimen length. Materials reinforced at an angle to the specimen's longitudinal axis fail in shear without crumpling at the end faces because all of the shear load is taken up by the matrix. The aforementioned basic modes of failure can be accompanied by a series of other phenomena: inelastic and non-linear deformation of the reinforcing fibers and matrix, delamination, surface peeling, overall buckling and crushing of the end faces. Failures with different combinations of these phenomena can make the determination of the failure mode very difficult. In compression testing, great care must be taken to ensure stability of the specimen, especially in the gage section. Buckling of the specimen side face is not always detectable and will cause erroneous strain measurements. Special test fixtures are used to prevent overall buckling of the specimen.
788 Mechanical tests 34.4.3 TENSION TESTING OF RINGS
The stress concentration problem is eliminated in tests that use uniform internal The most popular means of tension testing pressure generated by the use of a compliant rings uses a half-disk loading device (Table (d)) or by a hydraulic system ring (Method 34.2, Methods (a) and (b)). This is because the (Method (e)). The disadvantages of the comtest is easy to perform and the data is easy to pliant ring test method are the need for analyze. However, it has several significant periodic calibration of the loading element disadvantages: the strain distribution over the and the need for very careful preparation of specimen circumference is non-uniform, fricthe specimen surface. The disadvantage of the tion between the specimen and the half-disks hydraulic test technique is the need for expenhas a strong effect on the results, and there is a stress concentration in the specimen at the gap sive hydraulic equipment. between the two half-disks because of a variation in the radius of curvature of the specimen. In thin-walled rings, the stress concentration 34.4.4 COMPRESSION TESTING OF RINGS takes the form of a slight increase in radial ten- Radial compression of rings is accomplished by sile stresses that causes a minor effect on the external pressure (Table 34.2).The analog of the test results. However, in thicker specimens, half-disk tension test is the simplest compresthe resultant interlaminar shear stresses can be sion test method (Method (c)). The primary high enough to cause failure at lower loads difference is that in the compression test, it is than failure due to the circumferential stresses possible to reduce the stress concentration in alone. The error increases with increases in the the specimen at the split line. The best results relative specimen thickness t / R , the degree of are obtained with a semi-circular housing that anisotropy, and the ultimate strain of the mate- has a locking arrangement that prevents radial rial. Since this test method yields erroneously growth at the split-line (Fig. 34.2). There are low strength values, it can only be used for also compression analogs for the compliant qualitative comparison of composites. ring (Method (f)) and the hydraulic system Corrections have been suggested but have not (Method (g)).In the compliant ring method, the been used in practice. compliant ring is the elastic foundation of the
Fig. 34.2 Typical interlocking features for ring compression fixture.
Shear 789 specimen and to a certain extent it prevents buckling of the specimen. The external pressure may also be applied by mechanical devices such as multiple cam. The primary difficulty in compression testing of rings by external pressure is the selection of a relative thickness, t / R , to reduce secondary loading effects. Depending on the relative thickness of specimen, t / R , and the degree of anisotropy of the material E,/G,, three different failure modes have been observed. Thin-walled rings fail by buckling, thick-walled rings fail by biaxial compression, and optimum thickness rings fail by circumferential compression. Analysis of the test data must consider the radial as well as circumferential stresses. Delamination of the inner layers of the ring makes it difficult to correctly calculate the compressive strength. This delamination is often noisy. The occurrence of this failure mechanism also depends on the relative thickness of the ring. Delamination of helical windings can lead to unwinding of the specimen. Hoop wound rings can fail by layer-by-layer delamination. 34.5 SHEAR
34.5.1 IN-PLANE SHEAR
Shear properties, especially shear strengths, are difficult to measure. The simple and economical rail shear test is often used for this purpose (Table 34.3, Methods (a) and (b)).The extent of edge effects and the uniformity of the shear stress distribution over the specimen width depends on the length-to-width ratio of the specimen gage section L/w and on the relation of elastic constants G,/E of the material. Edge effects are negligible tor L/w > 10. Edge effects cannot be eliminated for materials with v, = vyxG -1. Thus, this method cannot be used for such materials. The elastic constants obtained by the rail shear test are less sensitive to the relative ratio, L/w, since the measurements are taken in the center of the specimen
gage section where the state of stress is the most uniform. However, edge effects have considerable influence on the shear strength. Therefore, it is better to bond the specimen to the rail links than to use mechanical fasteners. The stress distribution is not affected by the loading direction, i.e. along the diagonal or parallel to the sides of specimen gage section. Measuring the shear properties by tensile loading of an anisotropic strip is distinguished by its apparent simplicity (Table 34.1, Method (f)). The strip can have one of several different fiber lay-ups. This method is not used to determine in-plane shear strength because it yields low values. A state of pure shear is not assured even with a +45" lay-up. A similar test method involves tensile loading of a strip of a unidirectional material cut at an angle, 8, to the reinforcing fibers. The optimum angle is the one for which the relative shear strain y 1 2 / ~is x maximized and the shear stress rI2reaches its critical value. This angle depends on the anisotropy of the elastic and strength properties of the material tested. For advanced composites, the optimum angle is 10 to 15". Because the stress ratios are very sensitive to changes in the angle, rigid tolerances, 4 degree, are set on the specimen cut angle, the strain gage angle, and the direction of loading. In order to ensure that the stress state is uniform, relatively narrow strips, L/w = 14 to 16, are used. The in-plane shear modulus is often measured by twisting a square plate with four point loading (Method (c)). The wide acceptance of this method may be attributed to the simplicity of its calculations. However, the experiments should be performed with utmost care. This method is only applicable for small deflections, wp < O.lt, on plates made of materials which are uniform in thickness and orthotropic along the specimen axes. Test results for several different materials have shown that the P-wp relationship remains linear up to wp/t z 1. However, in practice, the deflection, w should be limited to 0.5t to preP' vent instability. Only the initial linear section
790 Mechanical tests of the P-wp curve should be used to determine the shear modulus. The optimum range of relative plate thickness, L / t , is determined by two conditions: the contribution of transverse shear to the deflection at small values of L / t and the possible loss of stability at large values of L / t . The limits of L / t are given for BFRP in Table 34.3. However, tests run on GFRP (glass fiber reinforced plastic), CFRP (carbon fiber reinforced plastic) and BFRP (boron fiber reinforced plastic) with different fiber lay-ups have shown that reliable data can be obtained at L / t > 15. The specimen must be flat and of constant thickness because the calculated shear modulus is related to t3. The distance from the point of support or load application to the corners of the plate should not exceed 2t. Experimental evaluation of these three shear test methods has shown that they all yield comparable values of in-plane shear modulus9. The successful application of the double Vnotch or Iosipescu shear test (Method (d)) to all types of fiber lay-ups is well known'O. It is invaluable for testing spatially reinforced composites since these materials do not possess planes of low shear strength. All other shear strength test methods use this characteristic of laminated composites to induce shear failure first and therefore are useless for spatially reinforced materials. In-depth investigations have shown that the Iosipescu method and its modifications yield good results in shear tests of carbon-carbon composites reinforced along three mutually perpendicular directions (3-D) and four principal diagonals of cube of 4-D". Spatially reinforced composites are less sensitive to the dimensions of the notches and gage length than laminated composites. The distribution of shear stresses is essentially uniform throughout the gage section of 3-D and 4-D materials cut at a 90" angle with a total notch depth equal to a half of the specimen thickness. Moreover, it is possible to prevent stress concentrations at the notch tips. Specimens with extra side notches yield the best shear strength data.
34.6 TORSION
Torsional loading of thin-walled tubes is a standard test for measuring in-plane shear modulus and strength (Table 34.4, Method (c)). In this test, the stresses are distributed uniformly around the circumference and along the length of the specimen. The shear strains are practically constant through the thickness of the specimen wall. In torsion, the definition of 'a thin-walled tube' is a function of the degree of material anisotropy EJE,, which can vary over a wide range. The disadvantages of this method are that it requires relatively large specimens, special test fixtures, inserts to prevent buckling of some specimens and wound specimens or specimens of special configurations, e.g. materials in which the fiber lay-up is parallel to the specimen axis. The results obtained by torsional shear tests compare favorably with results obtained by test methods using flat specimens. Torsional loading of split rings is also used to measure shear moduli (Method (b)). If the sample size limitations indicated in Table 34.4 are followed, bending effects are negligible. 34.6.1 INTERLAMINAR SHEAR
Good estimates of interlaminar shear properties, especially for spatially reinforced materials, have been obtained by torsion testing of rods with a circumferential notch (Method (a)).The specimens can be tested with or without a central bore. The important geometric parameters of the notch are the relative width, Lp/d, diameter, d, and wall thickness, t. It has been shown that within a range of L / d = 0.2-1.0, the length of the notch does not afPect the measured shear strength, tnu.The gage section diameter can be increased from 5 mm to 15 mm (0.6 in) without affecting rnu. However, increasing the diameter beyond 15 mm (0.6 in) causes a sharp drop in the measured strength.
Bending 791 34.7 BENDING
for measuring interlaminar shear strength. However, refined analysis has shown that the 34.7.1 THREE-POINT BENDING state of stress in a short bar of anisotropic The most popular type of bend test is the material is significantly different from the three-point bend test (Table 34.5, Methods (a) state of stress predicted by isotropic theory2J2. and (b)).The four-point and five-point bend The shear stresses through the thickness of a tests are less popular in spite of their consider- relatively short anisotropic bar have a paraable technical advantages over the three-point bolic distribution only in the middle of the span. At the loading points, the distribution of bend test. Theoretically, the moduli of elasticity of shear stresses through the thickness of the homogeneous materials in tension, compres- specimen has peaks near the surface directly sion, and bending are the same, i.e. Exf = E; = beneath the loading points. In relatively short E:. However, due to imperfections, the state of anisotropic bars, there are no planar regions of stress in bending and differences in the fiber constant maximum shear stress. Moreover, on lay-up through the thickness of the material, relatively short bars (L/w 4 5), the compresthe bending modulus of elasticity E: can differ sive transverse stresses from the load somewhat from Exf or E;. This difference is application points can extend over the entire length of the specimen and can exceed mean emphasized by the superscript ’b’. The formulae used to determine the elastic shear stress by up to a factor of 15. These combending constants, E: and GxZb, from the deflec- pressive stresses constrain crack opening at tion of a bar at its midspan must take into sample delamination and result in an apparent consideration the effect of interlaminar shear. increase of interlaminar shear strength. As a The effect of interlaminar shear can be result of these deviations from the ideal paraneglected for large values of relative span bolic stress distribution, the experimentally length L / t . For highly anisotropic materials, the determined interlaminar strength appears to decrease with increasing relative span width. relative span length must be greater than 40. Therefore, shear testing of relatively short bars When determining bending strengths, failcan provide only a qualitative comparison of and ””, ures limited by the normal strength, ., different composites. failures limited by the shear strength, zxF, Interlaminar shear strength can also be must be separated. Unlike those of isotropic materials, the two strengths for composite measured by three-point bending of curved materials can differ by an order of magnitude. segments. The shear strength is calculated The shear stresses can have a considerable with the same equations used for bending of prismatic bars. However, the additional intereffect on the failure. laminar normal stresses unique to curved Failure due to normal stresses occurs by beams must be taken into consideration. The fracture of the extreme outer layers in comnormal stresses act over the entire length of pression or tension. Failure due to shear the specimen. The sign of the stress depends stresses occurs by delamination approximately on the orientation of the specimen. In the case at the midplane of the specimen. Laminated of segments loaded with their convexity materials can fail by a violent debonding of the upwards (center load applied to the outer compressed outer layer. Very short bars experidiameter of the specimen) the stresses are tenence a third failure mode. They fail by 0 : . When the convexity is downwards, the sile, crumpling and shearing which is accompanied radial stresses are compressive, a;. In the forby an apparent increase in shear strength. Three-point bend testing of short bars or mer case, shear and tensile radial stresses ring segments is the most widely used method combine to decrease the apparent shear strength. In the latter case, the compressive
792 Mechanical tests
and 11),and the Arcan Test (mixed Modes I and 11). The geometry of the specimens, preparation methods, and analysis procedures have been described in detai113J4. Advanced composites, especially those reinforced with carbon and aramid fibers, have highly anisotropic thermophysical properties. This is reflected in the thermomechanical behavior of the structures fabricated with these materials. The combination of a polymer 34.7.2 BENDING OF RINGS matrix having a high coefficient of thermal Bending of complete rings by diametrically expansion and fibers having a negative coeffiopposed loads (Method (c)) is used to deter- cient of thermal expansion allows the mine elastic and strength properties of fabrication of composites with extremely low composites. Reliable results are obtained if the thermal expansion. This property of composrelative specimen thickness, t l r , is properly ites is used in the fabrication of structures selected. The acceptable range of relative which are stable over wide temperature thickness for determining the shear modulus, ranges. Low thermal expansion is commonly Go,: is based on the material anisotropy, achieved in two directions and the process can E,b/G,b. The shear modulus is calculated from be extended to materials which are spatially the load-diametral deflection data using the reinforced in three directions (3-D) or along same equations for three point bending of four diagonals of a cube (4-D)15.These comprismatic bars with a correction factor for the posites possess a thermal expansion coefficient fraction of deflection induced by shear which is both isotropic and very low. Measurement of the thermal expansion stresses. When the test is used to determine the interlaminar shear strength, rOpU,the rela- coefficients of carbon and aramid composites, tive specimen thickness must be chosen to especially in the reinforcement direction, is not ensure failure by shear delamination at the a trivial experiment. It must employ modem specimen mid-radius rather than failure by dilatometer and interferometer methodsI4. normal circumferential stresses, a?.
radial stresses impede the growth of the shear delamination crack and raise the apparent interlaminar shear strength. For accurate determination of the interlaminar shear strength, r&bu,segment dimensions must be selected so that the normal circumferential stresses, as, and normal radial stresses, or,are negligible compared to the shear stresses.
34.9 STRUCTURAL TESTING
34.8 SPECIAL TESTS
Laminates can fail by interlaminar delamination along specific planes. Toughness data is as important for characterization and failure prediction of composites as strength and stiffness data. Cracks in composites can propagate by Mode I, (crack faces opening normal to the crack plane), by Mode I1 (crack faces sliding in their planes), or by Mixed-Mode (combination of Mode I and Mode 11).The interlaminar fracture tests include the Double Cantilever Beam Test (Mode I), the Edge Delamination Test (Mode I), the End Notched Flexure (Mode 11), the Notched Three-Rail Shear Test (Mode 11), the Cracked Lap Shear Test (mixed Modes I
The unique challenges of designing and testing composite parts are due to the fact that the material and its micro- and macro-structures are created at the same time as the part. The design of critical structures must include the design of the material and must consider the unique behavior of composite materials which is influenced by processing techniques, actual service loading, and environmental conditions. Structural testing should start with tests of small-scale models fabricated by the same manufacturing process as the full-scale structure, followed by tests of prototype parts, specimens cut from structural elements, and finally, full-scale tests9J9.
References REFERENCES
1. Lubin, G. (ed.). Handbook of Composites, New York: Van Nostrand Reinhold, 1982. 2. Tamopol‘skii, Yu.M. and Kincis, T.Ya, Static Test Methods for Composites. New York: Van Nostrand Reinhold, 1985. 3. Witney, J.M., Daniel, I.M. and Pipes, R.B.,
793
11. Greszczuk L.B., Shear Modulus Determination of Isotropic and Composite Materials. ASTM
Special Technical Publication, 1969,460: 140-9. 12. Uemura M., Problems in Mechanical Testing Methods of Advanced Composite Materials. In Proc. 10th Tsukuba General Symp., 1990, pp. 43-54. 13. Pagano, N.J. (ed.), Interlaminar Response of Experimental Mechanics of Fiber Reinforced Composite Materials, (Composites Materials Composite Materials, Rev. Ed., Society for Series; 5) Amsterdam: Elsevier, 1989. Experimental Mechanics. Englewood Cliffs, NJ: 14. Carlsson, L.A. and Pipes, R.B., Experimental Prentice-Hall, 1984. Characterization of Advanced Composite Materials. 4. Kelly. A. (ed.). Concise Encyclopedia of Composite Englewood Cliffs, NJ: Prentice-Hall, 1987. Materials, Oxford: Pergamon Press, 1989. 15. Tarnopol’skii, Yu.M., Zhigun, I.G., and 5. Tarnopol’skii, Yu.M. and Vasiliev. V.V. (eds.). Polyakov, V.A., Spatially Reinforced Composites. Structural Composites. A Handbook, Moscow: Lancaster: Technomic, 1992. Mashinostroyenie, 1990. 16. Mechanical Testing of Advanced Fibre Composites, 6. Tsai, S.W., Theory of Composites Design, Dayton, Imperial College of Science and Technology, Paris and Tokyo: Think Composites, 1992. London: University of London, 1992. 7. W y o m i n g Test Fixtures. High Performance Test 17. Sims G.D., Nimmo, W., Johnson, A.F. and Fixtures. Product Catalog, Laramie: Wyoming Ferriss, D.H., Analysis of Plate-Twist Test for InTest Fixtures Inc, 1993. Plane Shear Modulus of Composite Materials. 8. Peters, S.T., Humphrey, W.D. and Foral, R.F., Teddington: National Physical Laboratory, 1992. Filament Winding Composite Structure Fabrication. 18. Lee, S. and Munro, M., Evaluation of in-plane Covina: SAMPE, 1991. shear test methods for advanced composite 9. Nikolaev, V.P., Panfilov, N.A., Popov, V.D., and materials by the decision analysis technique. Sinitsyn, E.N., Analysis of the Failure Composites, 1986,17(1), 13-22. Mechanism of Large-Scale Structures. Mech. 19. Nikolayev, V.P., Popov, V.D. and Sborovskii, Composite Mater., 1993,29(2):203-11. A.K., Strength and Reliability of Wound 10. Pinderra, M.J., Gurdal, Z.C., Hidde, J.S. and Fiberglass Reinforced Plastics. Leningrad: Herakovich, C.T., Mechanical and Thermal Mashinostroyeni ye, 1983. Characterization of Unidirectional Aramid/ Epoxy. Report CCMS-86-08, VPI-E-86-29, Virginia Polytechnic Institute and State University, Blacksburg, VA, 1987.
DURABILITY AND DAMAGE TOLERANCE OF FIBROUS COMPOSITE SYSTEMS
35
Ken Reifsnider
35.1 DEFINITIONS AND ISSUES
Durability and damage tolerance are critical to the design of composite structures. Damage tolerance is the approach often required for the certification of safety-rated structures such as aircraft components; durability has been identified as one of the most important technical drivers for the design of major composite structures such as the High Speed Civil Transport. Recent reports from the National Materials Advisory Board and a great volume of other literature focus on these Of course, there are many nuances in the definitions of durability and damage tolerance. However, the basic concepts are quite
simple, and are illustrated in Fig. 35.1. Damage tolerance is the remaining strength after some period of service, and durability, in general, has to do with how long the component will last, i.e. with the life of the structure. In this context, durability is often discussed in terms of the resistance or susceptibility to damage initiation. Both concepts imply that the subject component is being exposed to applied conditions such as mechanical loading and environments such as temperature and chemical agents over long periods of time that are typical of the projected service life of the component. There are several technical concepts that form a foundation for our discussion of these closely related topics. The first of these is the
Damage Tolerance (Remaining strength) 1
Normalized stress level Life Locus Durability (Life)
4
Time / Cycles Handbook of Composites. Edited by S.T.Peters. Published in 1998 by Chapman & Hall,London. ISBN 0 412 54.020 7
Fig. 35.1 Basic definitions of ’durability’ and ‘damage tolerance’.
Definitions and issues 795 question of the relationship of material strength to structural strength. In general, the strength of (fiber reinforced) composite materials is represented by an array of values that reflect the anisotropic nature of the materials (Fig. 35.2). For planar materials, at least the tensile strength and compressive strength in the fiber direction and transverse to the fibers and the in-plane shear strength are required for a complete answer to the question of ’how strong is this material’. However, as an array, those values do not directly show ’how strong is a composite structure’. Several possible answers to that question are typically given. One may use a ’failure criterion’ that compares all of the point stress components with all of the material strength components (such as the Tsai-Hill or Tsai-Wu riter ria)^ in some collective form based on concepts such as critical energy, critical shear resistance, etc. The salient point to be made is that the complexity of (inhomogeneous) composite materials and their array of anisotropic material strengths give rise to the development of a corresponding array of damage and failure modes in these materials that must be understood and correctly modeled to answer the question of
’how strong is this composite structure’, even if the array of material strengths are known (shown in Fig. 35.2). Hence, there is a need to develop understandings and representations of the critical damage and failure modes that control the performance of engineering components. This technology is currently incomplete, but discussions of those topics will follow. A second fundamental concept is microstructural architecture. As shown in Fig. 35.3(a), many fibrous composite components are made in layered or laminated form, with the fibers in different layers having different directions; in some cases the plies are made from different materials to form a ’hybrid’ composite. In addition, the fibers may not be straight, but may be woven, braided, or arranged in mats of various types (Fig. 35.3(b),(c)). These details have a major influence on the durability and damage tolerance of the materials. In fact, most composite material systems are ‘designed’ to be ’fiber dominated’, to take advantage of light, strong and stiff (but brittle) fiber materials that are available. Typically, the fibers, their geometry and their arrangement are important parts of the question.
Five in-plane strength values for fiberous composites: Tension and compression strength in the fiber direction Xt or X,
Tension and compression strength in the direction transverse to the fibers Yt orY,
in-plane shear strength
f 1
Composite
-S
Strength tensor:
IL-kr
*
Fig. 35.2 Schematic illustration of ’principal strength’ directions in a unidirectional continuous fiber composite laminate.
796 Durability and damage tolerance offibrous composite systems
Fig. 35.3 Typical engineering composite reinforcement types: (a) fibrous, unidirectional pile; (b) fibrous woven; (c) fibrous, braided.
A third technical issue has to do with the degradation of intrinsic strength and stiffness. For metals, the material stiffness and strength are generally constant during the life of the engineering component. This may not be true for composites. Stiffness changes of the order of 10-20% may be caused by micro-cracking, for example. Since many structures are stiffness designs, this mode of degradation must be considered. In addition, the intrinsic material strengths (indicated in Fig. 35.2) may also be degraded, especially by such things as physical or chemical aging. This behavior must also be part of the supporting predictive technology developed for these materials. Nondestructive methods of tracking such degradation are under development, but this remains as a challenge currently. Methodologies for the assessment and prediction of durability and damage tolerance of composite materials typically involve the following features:
0
0
0
0
Remaining strength and life models are developed and predictions are made for each independent failure mode (such as fiber failure in tension or micro-buckling in compression, etc.). Mechanics representations of the state of stress and state of material are constructed on the basis of a 'representative volume' of the material that is typical of the distributed damage state that controls the remaining stiffness and strength of the composite. A typical representative volume of material is a controlling ply in a laminate, but may be a micro-buckling ligament, a small group of fibers, etc. Various methods are used to characterize and monitor the rate of strength degradation in composites. A typical parameter which is useful for that purpose is stiffness change; however, that parameter is not appropriate in some cases. Micromechanics (mechanics analysis at the
Damage modes and failure modes 797 fiber/matrix level of representation) is increasingly used for remaining-strength modeling, for the calculation of stiffness change (which leads to internal stress redistribution), and for the estimation of remaining strength for a given failure mode. Statistical considerations are essential for the correct representation of the long-term behavior of composites. Composites typically fail because of the statistical accumulation of defects, which eventually interact to create a critical condition. This is in contrast to self-similar single crack propagation that is the typical mechanism of failure for common metals. Time-dependent behavior such as viscoelastic creep, creep rupture (driven by such things as internal stress redistribution or oxidation), and aging are typically important in the consideration of the longterm durability and damage tolerance of polymer composites, particularly for components that serve at elevated temperatures. This chapter will discuss the range of physical and engineering details that define and control this subject. Of course, a complete discussion would fill several volumes, so the reader should regard this discussion as only a starting point for further study.
35.2 DAMAGE MODES AND FAILURE MODES
The failure of 'typical' (homogeneous isotropic) engineering materials is a familiar topic. The subjects of ductile rupture and brittle fracture are widely discussed and taught in undergraduate and graduate courses. However, composite materials generally do not behave in a manner easily described by either plasticity (or yield) theory or by selfsimilar crack growth concepts. The reason for this different behavior is the fundamental difference in the micro-structure of composite materials, i.e. in the way they are
made. Composites consist of mechanical 'mixtures' of distinct phases (such as fibers or particles) in a matrix material. The geometry and arrangement of the reinforcement phase is carefully chosen to achieve the desired composite properties. As a result, such material systems are always inhomogeneous, often anisotropic, and often brittle. These three basic characteristics control the nature of damage development and failure in composite materials. The most salient single feature of damage in composites is the process of damage accumulation. Damage development usually involves many damage modes which create a widely distributed damage state, and failure is usually the result of a statistical accumulation of damage (rather than the statistical occurrence of damage). As discussed below, these multiple damage accumulations on failure modes are often closely related to the manner in which the composite is made, especially to the basic nature of the inhomogeneity and anisotropy of the material. This damage development process ultimately controls durability and damage tolerance, so we will discuss some typical major features of that phenomenon. The most pervasive damage mode in composite materials is microcracking, most often in the matrix material. Figure 35.4 shows two embodiments of this mode. Figure 35.4(a) shows an X-ray radiograph of a cross-ply laminate with cracks in both ply types, and Fig. 35.4(b) shows matrix cracking parallel to the fibers in the off-axis plies of a laminate, as seen from a tracing of those cracks as they appear on the edge of this [0,45,45,90Is laminate. A typical scenario for the development of such cracks is the formation of matrix cracks as a function of increasing applied load or increasing cycles of loading. These cracks typically extend through the thickness of a ply and generally extend quickly in the fiber direction if the local stress is uniform. Several other important features of matrix cracking are suggested by Fig. 35.4. As shown by Fig. 35.5, matrix crack formation releases
798 Durability and damage tolerance offibrous composite systems
"'r!
'r!-.eI
rt .1!
.I:
I
Strain Fig. 35.5 Change in slope of the elastic stress-strain curve induced by microcracking. APPLIED STRESS (MPo) 200 300 400 500
100
I
I
b
600 I
I
m
Fig. 35.4 Microcracking in the matrix, parallel to the fibers; a radiograph of a cross-ply laminate with (a) inter-ply delamination at crack intersections (arrow) and (b) a tracing of matrix cracks on the edge of a [0,+45,-45,90] laminate.
stored energy in the cracked ply or material, and changes the stiffness of material proportionately, a matter of concern to engineering applications, as noted earlier. However, the density of cracks in the ply of a laminate reaches a stable saturation value, as first observed by Reifsnider et a1.5,8,called the characteristic damage state (or CDS) of the ply. That CDS is a function only of the properties of the plies, their thickness, and their stacking sequence. Figure 35.6 shows that the same CDS is formed by static or cyclic loading. This CDS can be readily predicted since the crack spacing is determined by the rate at which the surrounding material can transfer stress back into the broken ply. Moreover, the stiffness change caused by this cracking can also be predicted as well7r9-*l. A second important damage mode is delamination, as shown in Fig. 35.7. Delamination is driven by the fact that local regions of the composite would deform differently in response to
01 0
I
I
I
I
0.2
0.4
0.6
0.8
NO.
I .o
OF CYCLES (MILLIONS)
Fig. 35.6 Data showing identity of the equilibrium crack spacing ('characteristic damage state' or CDS) for quasi-static and cyclic loading of a laminate.
the local loads if they were not bonded together in the composite. Hence, stored energy is released if they separate, and that energy drives the separation process. The most common example of this damage mode is the separation of the plies of a laminate near a free edge, as shown in Fig. 35.7. This process has been widely studied and is well described. More will be said of this driving mechanism below. It should be noted that delamination is usually nucleated by other damage modes (such as matrix and although it is a common damage mode, it is not usually a failure mode, per se. Delamination usually begins
Damage modes and failure modes 799 fW"82tr'tiKT W 7 K " V q T 4 ~ K ~ W m matrix ~ Tcracks in one ply may cause fiber fracture in an adjacent ply due to the local stress concentrations21,2z. Figure 35.8@) shows a second feature of importance. When the matrix and fibers have comparable stiffness and strength, the fibers may break many times along their length before the composite fractures. In this situation, fiber fracture can cause a significant I stiffness loss as well as a strength reductionu. Another generic damage mode is micro' buckling, induced by local or global compressive loads, as shown in Fig. 35.9.
-7
1
Fig. 35.7 Edge micrograph of delamination (arrow) showing (a) relationship to matrix cracking; (b) plan view radiograph of edge delamination in a crossply laminate (shaded regions).
at an edge, such as a cutout, bolt hole, rivet hole, etc. If it is in a region of nonuniform stress, it may stop growing when it reaches the boundary of that region. Even if it grows to large dimensions, it usually does not cause significant loss of strength in engineering sized structures. Still, the loss of integrity can lead to other damage and failure modes, so it should be avoided. A third generic damage mode is fiber fracture. Many composites are 'fiber dominated', i.e. they depend on the fibers for their stiffness and strength. Hence, fracture of the fibers is both an important damage mode and failure mode. However, fiber fracture is difficult to detect and has been studied less completely than many other damage modes. However, considerable data have been ~ollected'~'~. Figure 35.8 shows two examples of such data, driven by two important mechanisms. Figure 35.8(a) shows fibers broken beside one another, a typical situation. In many composites, the fibers are coated with a material that decreases the tendency for the fracture of one fiber to cause the fracture of neighboring fibers by forming an 'interphase region' around the fibers that tends to 'isolate' the fracture effectslS2O.It is also important to note that the
800 Durability and damage tolerance offibrous composite systems There are several aspects of this phenomenon that are of importance to durability and damage tolerance. For example, the compression strength (or remaining strength) of the composite may be controlled by the local stress required to initiate the local instability, in which case one wants a large diameter, stiff fibers in a stiff matrix. Or, the strength may be controlled by local resistance to shear deformation after buckling begins, in which case one would choose a tough matrix or interphase region between the fibers and the matrix. This is another case in which a damage mode may or may not be a failure mode, an important distinction.
Fig. 35.9 Localized microbuckling in a polymer matrix composite. Printed with permission, I.M. Daniel and 0. Ishai, Engineering Mechanics of Composite Materials, Oxford University Press, Oxford, 1994.
Another 'damage mode' of considerable importance to polymer composites and all composites used at high temperatures is the phenomenon of creep, i.e. time-dependent deformation at constant applied stress. Figure 35.10 shows a typical form of that behavior, with an initial transient region, a steady state region (in which most engineeringdesign is done), and a tertiary (usually unstable) region. This phenomenon is usually represented by introducing viscoelastic
Time
Fig. 35.10 Schematic of typical creep deformation at constant load and temperature.
or rheological models that represent the behavior in terms of a change in the stiffness of the material with time, as a function of temperature. Quite often, the reinforcing fibers do not show creep behavior at low temperatures, but at high temperatures, essentially all constituents may creep. The changes of stiffness with time can be characterized in the laboratory, and must be modeled carefully, based on those data. In fact, this part of the behavior is critically important to the correct calculation of internal stress states, since the creep of the constituents changes the internal stress distribution greatly in some cases. For example, if the matrix creeps more than the reinforcing fibers (a typical situation), that creep 'relaxes' the stress in the matrix, and increases the load carried by the fibers. If we wish to calculate a fiber-controlled strength, for example, a correct representation of this behavior must be included in our model. Finally, another failure mode is creep-rupture. This is a fairly general terminology used to refer to a variety of physical phenomena that produce time-dependent failure. This can be due to, say, oxidation of the fibers, or to other physical degradation processes which eventually cause rupture. It is clear that these phenomena must also be modeled correctly if we are to discuss durability and damage tolerance of material systems.
Damage drivers and damage 'resistance' 801 35.3 DAMAGE DRIVERS AND DAMAGE
'RESISTANCE'
In the previous section, a number of damage and failure modes that occur in composite materials, and ultimately control durability and damage tolerance, were identified. Many of these modes are related to the manner in which the composites were put together. This raises the basic question of 'can one design composite materials to be durable and damage tolerant?' Most of the rest of this discussion will address this question. Some general concepts will be followed by some micro-mechanics methods of quantifymg answers. Microcracking is likely to be the most pervasive damage mode in typical composites, especially under long-term loading, and most especially under cyclic loading. Even though most matrix materials are chosen because they offer some level of ductility, in most composite systems the matrix is highly constrained so that cracks develop due to local constraint, local stress concentrations, and local defects that grow rapidly under what is generally a 'plane strain' condition. Hence, matrix toughness, in the general sense, is the key to the reduction of matrix cracking. Increasing the strain to failure of the matrix material is a primary objective, and increasing the plane-strain fracture toughness of the matrix is a companion objective. There is a richly developed science and technology associated with matrix toughening; some starting points are listed in Wilkinson et al.24and Hedrick et al?5. A second damage mode identified earlier is delamination. This problem is a strong combination of structural and material concerns. The material concerns are essentially the same as those discussed for matrix cracking, with one important exception. Matrix toughness does not translate directly into interlaminar toughness. Hence, resistance to delamination cannot be controlled entirely with material property increases. The structural part of the problem does, however, present opportunities. It was mentioned before that delamination is driven
by local discontinuitiesin stress state, typically caused by neighboring plies or ply groups (bonded together) that would have very different strain states if they were not bonded. Hence, the orientation of the plies in a laminate and the stacking sequence of those plies are controlling players in the development of the interlaminar stresses that drive delamination. This problem has been exhaustively studied, and methods of reducing interlaminar stresses have been widely d i s c ~ s s e d ' ~ - ~ ~ , but because of the inhomogeneous and often anisotropic nature of composites, interlaminar stresses generally cannot be eliminated in laminated systems, so mechkical methods are widely used to control that tendency. The most successful of these is weaving, i.e. to use woven fiber architectures to reduce the anisotropy of a given ply, and therefore, to reduce the 'disagreement' between the response of any two or more plies. Woven materials are now widely used, especially for this reason. A second approach is to 'stitch' the composite in the region of non-uniform stress, typically near an edge of the laminate. Stitching simply 'clamps' the edge of the material to prevent it from separating; the internal stresses are still present. Stitching has a somewhat smaller number of proponents, but is a successful method as well. Finally, threedimensional reinforcement, such as mats or braids, also serve the purpose of providing constraint to the delamination drivers. These methods are not as widely used at this time, largely because of the difficulty associated with manufacturing. A less obvious influence on durability and damage tolerance is the bonding between the fiber and matrix. The nature of this influence has only come to light in recent years. Some of the mechanics models needed for this discussion will be developed in the next section; only a few general points will be made here. First, the properties of composite materials are determined not only by the properties of the constituents, but they are also greatly influenced by the manner in which the constituents
802 Durability and damage tolerance of fibrous composite systems interact. This critical interaction is, of course, controlled by the bonding between the constituents, between the fibers and matrix in our case. Typically, this bonding is ‘controlled’ by a fiber coating or ’sizing’. However, it is now known that such things as notched fatigue behavior can be improved by as much as two orders of magnitude by carefully ’designed’ ’interphase’ regions between the fibers and the matrixz6.There are at least two basic concepts operating in these effects. First, if one can toughen the composite by toughening the interphase between the fibers and matrix, the composite is likely to be more durable, as discussed above. Second, the interphase region can greatly influence the local stress state, and reduce the driving force for fiber-matrix debonding. An illustration of that is shown in Fig. 35.11. If one considers the strength of a composite under loads applied perpendicular to the fiber direction, then it is clear that the fiber causes a local stress concentration,in proportion to the difference between its properties and those of the matrix. However, if a coating around the fiber is introduced, this local concentration can be greatly reduced. In fact, for a ’rigid’ fiber, compared to the matrix, it is not surprising that a compliant coating on the fiber will increase the transverse composite strength by as much as a factor of two, and the
strain to failure by as much as a factor of In general, although design rules are not yet fixed, design of the interphase region is a new and important opportunity for the enhancement of the durability and damage tolerance of composite system^^^^*. The final subject in this section is ’failure criteria’; which are used to describe remaining strength. In general, failure criteria are chosen on the basis of the known failure mode. If fiber fracture controls strength, then a suitable criterion may be just the stress in the fiber direction divided by the strength in that direction. If matrix behavior is controlling, a shear stress or combined stress criterion may be appropriate. Figure 35.12 shows a comparison between strength ‘envelopes’predicted by two popular criteria. It is important to note that the inputs to the failure function will, in general, change as a function of time and loading history. The general form of any failure criterion will usually be some function of the ratios of stress in principal material directions to strength in those directions, as mentioned earlier. Under long-term conditions which induce damage, the local stress changes as damage causes redistribution, and the principal values of material strength change, due to such things as constituent degradation or micro-damage. Hence, to calculate damage tolerance by using
Interphase region Criterion:
I Maximum stress Applied
Stress (ksi)
0
90
Angle of Loading (deg)
Composite Fig. 35.11 Schematic diagram of the geometry of the interphase region in a fibrous composite, subiected to loading. ” transverse to the fibers.
Fig. 35.12 Allowable uniaxial loading as a function of angle of loading relative to the fiber direction in a unidirectional lamina, estimated from a maximum stress and a popular effective stress criterion.
Composite micro-strengthand remaining strength models 803 failure functions (or criteria) to calculate remaining strength, one must be careful to use the correct local stress state and material state in those expressions, especially when degradation has changed those states from their initial values.
utility of such models. The example is a recent model of tensile strength. (Figure 35.13). The stress in the broken fibers builds back up to the undisturbed level by shear transfer from the surrounding matrix, composite, and interphase region. That rate of buildup is directly proportional to the stress concentration in the next nearest fibers; if the buildup occurs over a 35.4 COMPOSITE MICRO-STRENGTH AND short distance (a short 'ineffective length), the REMAINING STRENGTH MODELS stress concentration in the neighboring fibers The importance of material principal strengths is great, and they tend to break causing very was noted, and the importance of composite brittle composite behavior. However, if the microstructure in the determination of those buildup occurs over a large distance (i.e. if the strengths has been emphasized. The proper- material around the fiber is very compliant or ties, geometry, arrangement, and bonding of breaks easily ), the strength of each fiber is lost the constituents determine the resulting val- completely when the first local fiber break ues of composite principal strengths. So, if occurs. A model has been developed that those factors are understood, strong, durable, describes the physics and mechanics of this damage tolerant composites can be designed. behavior, which estimates the fiber strength as: That understanding is currently incomplete, 2z0L l / m + l 2 l / m + l m + 1 but some models are available. Such models 4= ( K T ) m +2 are very valuable since they can tell us the preferable way to make composite materials, (1 + m)l/" (35.1) in contrast to how they can be made (the job of (C," + q m - 1 + ... + ly" the materials science community). In this limited space, one example will suf- where a, is the Weibull characteristic strength fice to demonstrate the general nature and of the fibers, z, is the shear stress between the
...'-+1(7r)
Composite
FlbersC
Fiber breaks
4 zt:
EE t
Normal stress In:
:
.
broken fiber
P
nelghboring
I
Average global values away from fiber fracture
Fig. 35.13 Schematic diagram of the local stress distribution around broken fibers in a unidirectional composite.
804 Durability and damage tolerance offibrous composite systems fibers and the matrix (usually taken as the continuous fiber reinforced composites is interphase strength), m is the Weibull shape outlined. A great many details will have to be parameter for the fiber strength distribution, omitted due to space limitations; the interand Cnis the local stress concentration when n ested reader can find them in other fibers are broken together in a local region. publication^"^^. Hence, the tensile strength in the fiber direcA start is to identdy a well-defined failure tion can be estimated on the basis of the mode, as defined earlier. Since damage is disproperties of the constituents and the inter- tributed, this damage mode will be 'typical' of phase region between the fibers and the any 'representative volume' of material; a matrix. If any of those constituent characteris- mechanics boundary value problem on such a tics change, the model can show how the representative volume (RV), as suggested in strength of the composite changes, i.e. the Fig. 35.14 can be 'set'. This RV may be disconmodel can be used to calculate the damage tol- tinuous; i.e. it may have cracks, delamination, erance of the composite if the failure mode is debonds, etc. But some part of it will remain controlled by fiber strength in tension. intact until fracture of the composite, and this Comparable models can be constructed for part of the RV that defines the fracture event is compression failure, and for other failure a 'critical element'. Therefore the objective is mode^^*^^. the calculation of the state of stress and state of material in the 'critical element.' One can write all failure functions, Fa, in that critical element, 35.5 ESTIMATION OF REMAINING and claim that when these failure functions STRENGTH AND LIFE (for each distinct failure mode) predict failure, As indicated earlier, damage tolerance is the composite will fail. defined by remaining strength, and durabilInvoking kinetic theory we can derive an ity is usually discussed in terms of life. In this equation that relates changes in stress state and final section, one approach to the estimation material state with time and loading history to of the durability and damage tolerance of remaining strength, i.e. allow the incorporation
failure modes
Fig. 35.14 Diagram illustrating how experimental observation of failure modes define the representative volume (used to set the boundary value problem) and the critical elements in which all continuum states are defined.
Estimation of remaining strength and life 805 of the explicit time, cycles, and environmental dependence that leads to phenomenological behavior such as creep, creep rupture, fatigue, and aging into the calculation of remaining strength. From thermodynamic principles, the following expression can be derived:
F, = 1 -
lyl(l
- Fa),(
N) n d( $) 1-1
sile fiber failure, and it is assumed that some fatigue behavior of unidirectional material under uniaxial stress in the fiber direction has been measured, a 1-D SN relationship can be derived, of the form:
s, = A + B (log S"
(35.2)
N)p
(35.3)
where, for our example, A = 1, B = -0.1, Su = 100 ksi (the initial ultimate strength), p = 1,and Sa is the applied stress amplitude. Equation (35.3) provides an input, N, into (35.2) since
where in the critical element, F, is the normalized remaining strength, n is cycles, and N is the life of the critical element under the current state of stress and state of material. The methodology of this calculation is shown in Fig. 35.15. Remaining composite strength, F , is calculated directly; life is calculated by the coincidence of Fa and Fr. Numerous comparisons of such calculations with experimental data have been made over the last 10 years or so, and there are a few examples at the end of this chapter. The immediate purposes are served by using eqn(35.2) to examine the effects of some hypothetical change in material state and stress state on remaining strength (damage tolerance). If the failure mode is ten-
State of
where u,, is now the current local ply stress in the fiber direction, and X , is the current local principal material strength in tension, given by eqn (35.1). Substituting eqn (35.4) in eqn (35.2) and assuming that no other phenomena are present (and using j = 1.2, a known typical value), the curve (a) in Fig. 35.16 results. Now, suppose the ply is the critical element in a multiaxial laminate having off-axis plies that crack and 'dump' stress into the critical element as a function of
State of
I
life
Subcritical
Critical
N,
I
N2
reipresentative volume
Fig. 35.15 Schematic flow diagram of the manner in which the MRLife simulation scheme calculates
remaining strength and life.
806 Durability and damage tolerance offibrous composite systems
I
Remaining Strength
-
I
Cycles
Cycles
Fig. 35.16 Calculated remaining strength predictions for (a)0" lamina degradation alone; (b) added
Fig. 35.18 Assumed degradation of fiber strength for the sample laminate.
effect of matrix cracking; (c) added effect of fiber degradation (e.g.by oxidation). cycles, according to the rate shown in Fig. 35.17 (from cracking rates that must be measured or estimated). With this internal stress redistribution, only, added to eqn (35.2), the damage tolerance changes to curve (b) in Fig. 35.16. Of course, if creep occurs in the matrix (perhaps because of increased temperature), in which case the local fiber stress will increase again as a function of cycles to change the form shown in Fig. 35.17. Finally, suppose creep rupture is occurring, driven, for example, by oxidation of the fibers that is reducing the diameter of the fibers, D, in eqn (35.1), in the manner shown in
FiberDirection Stress
40
Stress increase due
'
I 1
4
I
4
5x10 9.999~10 Cycles
Fig. 35.17 Assumed increase in stress in the 0" ply
due to matrix cracking.
Fig. 35.18. Then the strength model, eqn (35.1), correctly integrates that micro-change into the global calculation, and eqn (35.2) shows the damage tolerance to be curve (c) in Fig. 35.16 for that situation. Hence this 'micro-kinetic' approach has the capability to estimate durability and damage tolerance for very complex situations involving combinations of many time and cycle dependent phenomena in composite systems, using a mechanistic approach. An example follows. Using the methods described above, the rate of matrix cracking and the unidirectional SN curve of a carbon fiber reinforced PEEK matrix composite were determined, and used to estimate the remaining strength and life of several different laminates made from that material. Figure 35.19 shows an example of the predicted and observed life for several load levels of a quasiisotropic laminate made from such material and Fig. 35.20 shows comparisons of the predicted and observed remaining strength of such laminates for two load levels and cycles of load application. It can be seen that this approach can produce quite useable results. Many such predictions have been compared using the MRLife performance simulation code based on this
Estimation of remaining strength and life 807
AS-4lPEE K (APC-2) Quasi-Isotropic Notched Fatigue (R=-1) 0.75
0.45
3
I
1
I
4
5
6
7
Log N (Cycles) Simonds B Stinchcomb MRLife (1 989) Prediction 0
Fig. 35.19 Predicted (line) and observed life for a quasi-isotropicAS-4/PEEK notched coupon under fully reversed loading.
Residual Strength at 0.70 Suit
Residual Strength at 0.90 SI,,
11.05
1.05
1.w
1.00
r
r
P E!
F E!
5
5
3
3
0.95
; 2 .-
0.w
d ......................................................
x1 N
0.w
€0 z
0
z 0.85
0.05
P
P
.....................................................
0.85
......................................................
0.80
0.80
2
Cycles M E h Expepnt
5
1 0 2 0
5 0 1 0 0 2 0 0
Cycles Mfih Expepml
Fig. 35.20 Predicted (lines) and observed residual strength of AS-4/PEEK specimens subjected to under fully reversed cyclic loading.
808 Durability and damage tolerance offibrous composite systems 35.6 SUMMARY
This has been a short outline of the physical behavior associated with the durability and damage tolerance of composite material systems, and a few modeling approaches to the estimation and prediction of that behavior have been indicated. It should be noted that there is every evidence that composite materials are remarkably durable and damage tolerant. Fatigue allowables for carbon fiber reinforced polymer composites, for example, exceed those of structural steels, and the durability and damage tolerance of ceramic composites make them the only choice for ultra-high temperature applications in turbines, etc. Understanding of this subject, which is admittedly incomplete, has reached a level that is sufficient to support engineering applications of composites to even the most demanding situations in the most severe environments. In fact, that is exactly the situation in which the application of composites is most beneficial and cost effective. Composite material systems can provide many new opportunities to design for damage tolerance and durability. REFERENCES 1. Life Prediction Methodologies for Composite Materials, NMAB-460. Washington, D.C.: National Academy Press, 1990. 2. High-Temperature Materials for Advanced Technological Applications, NMAB-450. Washington, D.C.: National Academy Press, 1988. 3. Horton, P.E. and Whitehead, R., Damage Tolerance of Composites, Vol. I and 11, Air Force Wright Aeronautical Laboratories, AFWAL TR87-3030, 1988. 4. Daniel, I.M. and Ishai, O., Engineering Mechanics of Composite Materials. New York: Oxford Univ. Press, 1994. 5. Reifsnider, K.L. and Highsmith, A.L., Characteristic damage states: A new approach to representing fatigue damage in composite laminates. In Materials: Experimentation and Design in Fatigue, Guildford, U.K.: Butterworth/IPC, 1981, pp. 246-260.
6. Damage in Composite Materials: Basic Mechanisms, Accumulation, Tolerance, and Characterization, STP 775, American Society for Testing and Materials, (ed. K.L. Reifsnider), 1982. 7. Highsmith, A.L., Stijfrzess Reduction Resulting from Transverse Cracking in Fiber-Reinforced Composite Laminates. Master of Science Thesis. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1981. 8. Reifsnider, K.L., Some fundamental aspects of the fatigue and fracture of composite materials. In Proc. 14th Ann. SOC. Engng Science Mg. Bethlehem, PA: Lehigh University, 1977. 9. Highsmith, A.L., Damage Induced Stress Redistribution in Composite Laminates. PhD Dissertation. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1984. 10. Bader, M.G., Bailey, J.E., Curtis, P.T. and Parviz, A., The Mechanisms of Initiation and development of damage in multi-axial fibre-reinforced plastic laminates. In Proc. 3rd Intern. Conf. Mechanical Behavior of Materials. Cambridge, U.K., 1979. 11. Reifsnider, K.L., Damage and damage mechanics. In Fatigue of Composite Materials, (ed. K.L. Reifsnider), Amsterdam: Elsevier Science Publishers, 1990. 12. OBrien, T.K., Characterization of delamination onset and growth in a composite laminate. In Damage in Composite Materials, ASTM STP 775, 1982, p. 140. 13. OBrien, T.K., Analysis of local delaminations and their influence on composite laminate behavior. In Delamination and Debonding of Materials, ASTM STP 876, 1985, pp. 282-297. 14. OBrien, T.K. and Hooper, S.J., Local delamination in laminates with angle ply matrix cracks: Part I, Tension tests and stress analysis. In N A S A TM 104055,1991. 15. Razvan, A. and Reifsnider, K.L., Fiber fracture and strength degradation in unidirectional graphite/epoxy composite materials. In Theoretical and Applied Fracture Mechanics, 1991, 16,81-89. 16. Razvan, A., Bakis, C.E. and Reifsnider, K.L., SEM Investigation of fiber fracture in composite laminates. In Materials Characterization, 1990,24, 179-190. 17. Lorenzo, L. and Hahn, H.T., Fatigue failure mechanisms in unidirectional composites. In Composite Materials: Fatigue and Fracture, ASTM STP 907, American Society for Testing and Materials, Philadelphia, PA, 1986, pp. 210-232.
References 809 18. Ishida, H. (ed.), Controlled Interphases in Composite Materials, New York: Elsevier, 1990. 19. Warren, R. (ed.), Ceramic Matrix Composites, New York: Chapman and Hall, 1992. 20. Reifsnider, K.L. (ed.), Fatigue of Composite Materials, London: Elsevier Science Publishers, 1991. 21. Jamison, R.D., Fiber fracture in composite laminates. In Proc. lntl. Con$ on Composite Materials VI, 1987, no. 3, pp. 185-199. 22. Jamison, R.D., Schulte, K., Reifsnider, K.L. and Stinchcomb, W.W., Characterization and analysis of damage mechanisms in tension-tension fatigue of Graphite/Epoxy Laminates. In Efects of Defects in Composite Materials, ASTM STP 836, American Society For Testing and Materials, Philadelphia, PA, 1984, pp. 21-55. 23. Tiwari, A., The Development of an Interpretive
Methodology for the Application of Real-Time Acousto-Ultrasonic NDE Techniques for Monitoring Damage in Ceramic Composites Under Dynamic Loads. PhD Dissertation. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1993. 24. Wilkinson, S.P., Liptak, S.C., Lesko, J.J., Dillard, D.A., Morton, J., McGrath, J.E. and Ward, T.C., Toughened bismaleimides and their carbon fiber composites for fiber-matrix interphase Studies. In Proc. 6th Japan-U.S. Conf Composite Materials, 1992. 25. Hedrick, J., Patel, N.M. and McGrath, J.E., Toughening of epoxy resin networks with functionalized engineering thermoplastics. In ACS Advances in Chemistry Series, no. 233, Toughened Plastics I: Science and Engineering, (eds. C.K. Riew and A.J. Kinloch), 1993, pp. 293-304. 26. Swain, R.E., Reifsnider, K.L., Jayaraman, K. and El-Zein, M., Interface/interphase concepts in composite material systems. J. Thermoplastic Comp. Mater., 1990, 3, 13-23. 27. Case, S.W., Micromechanics of Strength-Related Phenomena in Composite Materials. MS Thesis. Blacksburg, Virginia: Virginia Polytechnic Institute and State University, 1993. 28. Carman, G.P., Eskandari, S. and Case, S.W., Analytical investigation of fiber coating effects on shear and compression strength, symposium on durability and damage tolerance, ASME WAM, (in press), 1994. 29. Jayaraman, K. and Reifsnider, K.L., The interphase in unidirectional fiber-reinforcedepoxies: effect of residual thermal stresses. Comp. Sci. Tech., 1993,47, 119-129.
30. Jayaraman, K., Gao, Z . and Reifsnider, K.L., The interphase in unidirectional fiber reinforced epoxies: effect on local stress fields. J. Comp. Tech. Res., 1994,16(1):21-31. 31. Jayaraman, K., Reifsnider, K.L. and Swain, R.E., Elastic and thermal effects in the interphase: Part 11. comments on modeling studies. J. Comp. Tech. Res., 1993,15(1):14-22. 32. Jayaraman, K., Reifsnider, K.L. and Swain, R.E., Elastic and thermal effects in the interphase: Part I. comments on characterization methods. J. Comp. Tech. Res., 1993,15(1):3-13. 33. Gao, Z . and Reifsnider, K.L., Micromechanics of Tensile Strength in Composite Systems. Fourth Volume, ASTM STP 2256, (eds W. W. Stinchcomb and N. E. Ashbaugh), Philadelphia, PA: American Society for Testing and Materials, 1993, pp. 453-470. 34. Xu, Y. and Reifsnider, K.L., Micromechanical modeling of composite compressive strength. J. Comp. Mater., 1993, 27(6):572-587. 35. Reifsnider, K. L. and Gao, Z., Micromechanical concepts for the estimation of property evolution and remaining life. In Proc. Intern. Con5 Spacecraft Structures and Mechanical Testing, Noordwijk, the Netherlands, 1991, pp. 653-657. 36. Curtin, W.A., Theory of mechanical properties of ceramic-matrix composites. J. Amer. Ceram. SOC.,1991, 74(11),2837-2845. 37. Reifsnider, K.L., Performance simulation of polymer-based composite systems. In Durability of Polymer-Based Composite Systems for Structural Applications, (eds A.H. Cardon and G. Verchery), New York: Elsevier Applied Science, 1991, pp. 3-26. 38. Reifsnider, K.L. and Stinchcomb, W.W., A critical element model of the residual strength and life of fatigue-loaded composites coupons. In Composite Materials: Fatigue and Fracture, ASTM STP 907, (ed. H.T. Hahn), Philadelphia, PA: American Society for Testing and Materials, 1986, pp. 298-313. 39. Reifsnider, K.L., Use of mechanistic life prediction methods for the design of damage tolerant composite material systems. In ASTM STP 2257, (eds M.R. Mitchell and 0. Buck), Philadelphia, PA: American Society for Testing and Materials, 1992, pp. 205-223. 40. Reifsnider, K.L., Evolution concepts for microstructure property interactions in composite systems. In Proc. IUTAM Conf. Microstructure-Property Interactions in Composite Materials. Aalborg, Denmark, 1994.
ENVIRONMENTAL EFFECTS ON COMPOSITES 36 Ann F. Whitaker, Miria M . Finckenor, H a r y W. Dursch, R.C. Tennyson and Philip R. Young
stabilizers, vertical fins and fairings were flight-tested with annual inspections. Composite usage has increased dramatically Satisfactory performance of the composite over the last three decades due to the advanmaterials was noted over fourteen years, with tages of light weight, specific strength and some parts experiencing more than 39 000 stiffness, dimensional stability, tailorability of hours of flight loads. Also evaluated were properties such as coefficient of thermal composite parts from military aircraft such as expansion and high thermal conductivity. the C-130 center wing box, S-76 tail rotors and Environmental effects on these properties may horizontal stabilizer, 206L fairing, doors, and compromise a structure and must be considvertical fin and the CH-53 cargo ramp skin. ered during the design process. Boron/epoxy, graphite/epoxy, Kevlar/epoxy Because of the variety of uses, the composand Nomex honeycomb were used in these ite environment cannot be exactly defined. aircraft and helicopter components. This chapter details the major environmental concerns for the composite designer, problems encountered with these environments in the 36.3 ENVIRONMENTS AND EFFECTS past, and some materials or protective systems effectively used. The use of trade names, how- 36.3.1 BIOLOGICAL ATTACK ever, does not constitute endorsement, either Biological attack on composites may consist of expressed or implied, by the authors. fungal growth or marine fouling. As reported in the literature, fungal growth does not 36.2 HISTORICAL PERSPECTIVES appear to be as damaging as the wet conditions that promote growth. Fungicide has been Use of composites in commercial aircraft mixed in with resins to retard this growth. increased under two NASA programs, the Marine boring organisms do not appear to Flight Service Evaluation Program and the attack glass-reinforced composites. Even Aircraft Energy Efficiency Program, begun in though marine organisms will grow on comthe 1970s. These programs included evaluaposite surfaces, mechanical properties do not tion of environmental effects on the composite appear to be affected, and the fouling can be parts of the Boeing 727, McDonnell Douglas by scraping (Fried, 1969). removed DC-10 and Lockheed L-1011 commercial airComposites with graphite fibers have been craft. Elevators, rudders, ailerons, horizontal used in medical applications for both internal and external purposes. Internal composite Handbook of Composites. Edited by S.T. Peters. Published structures, such as artificial joints or plates for in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7 bone fracture support, must be biocompatible 36.1 INTRODUCTION
Environments and effects 811 or the material may degrade over time. External composite designs, such as artificial limbs or orthotic braces, may experience impact damage, flexural and torsional loading during use. 36.3.2 FATIGUE
Fatigue, either through mechanical loads or acoustic vibrations, can cause crack growth or local defect formation. Fatigue design depends not only on the load, but also on the use temperature range and amount of moisture present. Very cold temperatures (below -50°C (-58°F)) may increase the stiffness of some composite materials, thereby increasing the susceptibility to fatigue damage (Staunton, 1982). Destructive effects of fatigue vary with the composite system tested. One example of fatigue resistance is the B-1 horizontal stabilizer torque box, an all-mechanically fastened hybrid composite structure (Staunton, 1982). Acoustic fatigue testing produced no degradation, nor did the service environment of moisture, mechanical fatigue, and temperature cycling from -12 to +167"C (10 to 260°F). 36.3.3 FLUIDS
Moisture Moisture effects on composites have been studied for decades. Water acts as a plasticizer when absorbed by the matrix, softening the material and reducing some properties of the laminate. Moisture may also migrate along the fiber-matrix interface, affecting the adhesion. Moisture in composites reduces matrix-dominated properties, such as transverse strength, fracture toughness and impact resistance. Lowering of the glass transition temperature may also occur in epoxy and polyimide resins with an increase in absorbed moisture. Debonding can occur due to formation of discontinuous bubbles and cracking in the matrix. Mechanical properties can be reduced even
further if heat is present or if the composite is undercured or has a large amount of voids. Moisture is absorbed into the composite until a saturation point is reached. This has been described as a non-Fickian process, meaning the rate of relaxation in the material due to water absorption is comparable to the diffusion rate of water. As the material properties change, such as a decrease in glass transition temperature, the diffusion process changes. Swelling stresses due to non-uniform water absorption have been investigated (Ashbee, 1989). Volume expansion due to water absorption can be a few percent at saturation. Moisture absorption is usually dependent on the matrix, but aramid fibers will also absorb water. The mechanical properties degrade in relation to the amount of moisture absorbed, with no further deterioration after saturation is reached. Strength reductions in polyester laminates have been found to be 10-1570, while epoxy resins are less vulnerable. In a few cases, drying of the composite restored the original mechanical properties. Testing of a glass/polyester laminate allowed to dry after ocean exposure at 1700 m (5700 ft) below sea level for three years showed little change in compressive strength and modulus, flexural strength and modulus, or interlaminar shear strength (Fried, 1969). Fiberglass composites with either polyester or epoxy resins have been used extensively in marine structural applications because of their strength-to-weight characteristics and resistance to the marine environment. Glass reinforcement is preferred over carbon fibers due to carbon's electrical conductivity, which may result in severe dissimilar metals galvanic corrosion with sea water acting as an electrolyte. MIL-HDBK-l7B, besides providing guidelines for characterizing materials and designing a composite system, contains a wealth of mechanical property and environmental effects data. The effect of moisture absorption or water immersion on weight,
812 Environmental eflects on composites (150°C (302°F)) environment, microcracking occurred. The amount of moisture absorbed, as measured by weight gain, is directly related to the change in mechanical properties. Salt water, antifreeze and gasoline had the most proAircraft fluids nounced effect. Dqmg did restore some, but The aircraft fluid environment consists of fuel, not all, of the strength and modulus. Volkswagen of Germany tested composite hydraulic fluid, lubricants, deicing comsystems for compatibility with gasoline, oil pounds, and water. Polysulfone has been and coolant for engine use (Beckmann and found to be sensitive to phosphate ester based Oetting, 1985). These materials were tested hydraulic fluids. Some polymer resins, such both as pure resin and in fiber reinforced lamas PEEK, may have lower glass transition inate form. Among the materials tested were temperatures after exposure to fluids with a glass, carbon and aramid fibers, high-temperhigh aromatic content. A study of stressed and ature epoxy, polyimide and polyester resin unstressed composite materials (Dexter, 1987) systems. Mechanical properties of the samples evaluated short-beam shear strength and tenwere measured after each 100 h of immersion sile strength after immersion in JP-4 fuel, 1000 h total. Adverse reactions of the carup to hydraulic fluid, a fuel-water mixture and bon reinforced materials with metal parts, oil fuel/air cycling for 5 years. The composites and combustion residues were noted. tested were T300/5208 graphite/epoxy, T300/5209 graphite/epoxy and Kevlar 49/5209 in (k45)s configuration. The Other fluids fuel-water immersion appeared to be the most damaging, reducing the tensile strength Liquids accidentally spilled on composite surof the T300/5209 and the Kevlar/5209 by 11% faces may also affect the mechanical and 25%, respectively. Fuel-water exposure properties. Methylene chloride, found in paint also degraded short-beam shear strength by strippers, may cause severe damage to epoxy resins and a number of other polymers. as much as 40%. Graphite/polyimide composites samples were immersed in various fluids for 10 min then Automotive fluids tested for flexure strength and modulus The automotive fluid environment consists of (Lisagor, 1979).Slight increases in these propgasoline, oil, battery acid, brake fluid, transmis- erties were noted for samples immersed in sion fluid and coolant. A study by University of hydraulic fluid, nitrogen tetroxide liquid, Michigan and General Motors (Springer, monomethyl hydrazine liquid and unsymmetSanders and Tung, 1981) details the effects of rical dimethyl hydrazine. Samples exposed to prolonged exposure of E-glass/polyester and hot hydrazine vapors were degraded beyond E-glass/vinylester to automotive fluids. the ability for mechanical testing. Solvents, Composite samples were immersed in water, bases and weak acids at room temperature do salt water, No. 2 diesel fuel, lubricating oil, not appear to affect graphite/epoxies and antifreeze or gasoline. Property changes mea- Kevlar/epoxies. Molten metal, such as alusured were weight, ultimate tensile strength, minum or titanium, may react with carbon tensile modulus, short-beam shear strength and fibers. shear (flexural) modulus. Specific materials exposed were OCF-E-920-1 polyester/E-glass, OCF-E-980 polyester/E-glass, and vinyl ester/E-glass. In a moist high-temperature
operational temperatures, and mechanical properties such as stress rupture characteristics is discussed.
Environments and efects 813 36.3.4 WEATHERING
effects on composite materials. Over 35 different types of organic matrix composites were Warm, moist climates may affect the perforflown on LDEF during its 69-month mission in mance of composites. Decreases of 10-20% in LEO. The post-flight testing and analysis of tensile strength have been noted for fibercomposites flown on LDEF has become the glass/polyester and fiberglass/epoxy (Graner, basis of understanding the long-term effects of 1982)where the surface resin has been eroded the LEO environment on composites. Data away due to extended weathering. Resins from the Solar Array Materials Passive LDEF which are more weather-resistant have been Experiment (SAMPLE), the Space developed since this study. In a 10-year study Environment Effects on Spacecraft Materials of real-time weathering of graphite/epoxy, graphite/polysulfone, and Kevlar/epoxy Experiment, the University of Toronto (Dexter, 1987), the materials that absorbed the Institute for Aerospace Studies (UTIAS) most moisture were most affected by UV radi- Experiment, and the Space Exposure of Composite Materials for Large Space ation. Erosion due to rain, snow or ice impact Structures Experiment are presented here and may be a problem for some aircraft parts, in the section on design considerations. such as radomes or leading edge parts. Further information may be found in the Coatings, such as polyurethane, may be used LDEF Post-Retrieval Symposium Proceedings to make composite parts more resistant to (Levine 1991,1992). this erosion. A real-time weathering study was performed by Grumman on fiberglass Atomic oxygen parts from E-2A and A-6A aircraft (Staunton, 1982). Length of service varied from 12 to 19 Atomic oxygen is formed through the dissociyears. Effects of weathering were dependent ation of 0, by UV radiation. It is the on the material used and whether a protec- predominant molecular species at 100-1000 tive coating was intact. When a hygroscopic km orbital altitudes. Its destructiveness is BF3400 curing agent was used, the fiber- caused by its strong chemical reactivity combined with its translational energy of 5 eV glass/Epon 828 epoxy lost nearly half of its J) from the high velocity of the spaceflexural strength. Fiberglass/Epon 828 epoxy (8 x craft. Studies on the effects of atomic oxygen with methylene dianiline/benzyl dimethyl have been performed both on samples amine (MNA/BDMA) curing agent performed well, retaining tensile and flexural exposed to the low earth orbit environment and in ground-based simulators. strength. Tensile and flexural test samples The Mass Spectrometer Incoherent Scatter taken from a fiberglass rotodome demon(MSIS) neutral atmosphere model is generally strated the value of good coatings. Where the used for predicting the atomic oxygen fluence paint was intact, the material retained more during a mission. Orbital altitude, inclination, than 90% of its original strength and 82-94% other orbit parameters and solar activity are of modulus. Where the paint had been eroded used as inputs. The amount of atomic oxygen away, the composite retained only 68% of its received by a surface also depends on its orioriginal strength. entation to the RAM or velocity vector. RAM facing composites (facing into the velocity vector) flown on LDEF were sub36.3.5 SPACE jected to an atomic oxygen fluence of The Long Duration Exposure Facility (LDEF) approximately 9 x IOz1 atoms/cm2 which has provided a wealth of information on the resulted in a thickness loss up to 0.013cm low-earth-orbit (LEO) space environment (0.005 in) of material, the equivalent of
814 Environmental effects on composites I
3
Meteoroiddebris impacts
A spacecraft in any orbit is susceptible to micrometeoroid impact. These small particles, fragments of asteroids or comets, may impact at velocities up to 60 km/s (37 mi/s) but average 17km/s (10.5mi/s). Those spacecraft in near-earth orbit are also susceptible to impact from pieces of space junk or debris, also travelling at high speeds. Damage from impact may reduce the strength of composite structures or cause rupture in filament-wound tanks. Damage may consist of cratering, penetration, including penetration of thermal or protective coatings, and spallation Fig. 36.2 is Fig. 36.1 Cross-sectional photomicrograph of of a classic Whipple-type meteoroid/debris atomic oxygen exposed LDEF graphite-reinforced shield before and after particle impact. OMC showing approx. one ply (0.013 cm/O.OOS in) of erosion.
e-
approximately one ply of laminate (Fig. 36.1). Impact particle For unidirectional reinforced specimens, the reduction in mechanical properties was proportional to the reduction in cross-sectional area. The SAMPLE experiment flown on the / One or more protectivetpmpers LDEF contained tensile specimens of graphite/epoxy systems and S-glass/epoxy. The composite systems flown were HMF Pressure wall 322/P1700 polysulfone in a ( d 5 ) s weave, HMS/934 in both 0" and 90" unidirectional Before imDact configurations, and P75S/934, also in both 0"and 90" unidirectional lay-ups. The Sglass/epoxy was flown with and without Debris cloud aluminized thermal control tape as a protective coating. These samples received direct atomic oxygen, as well as UV radiation, thermal cycling and micrometeoroid/debris impacts. Atomic oxygen reaction efficiency was calculated to be approximately 1 x cm3/ atom. The thickness loss due to atomic oxygen Penetration or erosion of the S-glass/epoxy samples in this 0 0 spallation of f pressure wall may experiment is estimated to be 9.14 pm (0.36 Spallation or may not occur mil). The glass fibers are not susceptible to erosion, and thus protect the underlying matrix. After imDact The atomic oxygen reaction efficiency for these samples was calculated to be 0.13 x lowz4 Fig. 36.2 Hypervelocity impact of Whipple bumper cm3/atom. design.
/
-
I
Environments and efecfs The potential hazard of a meteoroid or debris particle is dependent on its size, velocity, density and angle of impact. NASA SP-8042 may be used to predict the meteoroid environment encountered by a spacecraft in near-earth orbit, earth-to-moon space and near-lunar orbit. The space debris environment in earth orbit is continually changing as more debris is being added with every launch. At the time of publication, NASA TM-100471 is being used as the definition of the space debris environment. Prediction of damage caused by meteoroid/space debris impact can be made either by using a variety of models, study of impact sites in flight hardware, or simulation in the laboratory. Hydrocodes such as CTH and HULL use finite elements or finite differences to predict penetration or spallation of a composite laminate. Inspection of composite samples from LDEF (Fig. 36.3) revealed 2-5 impacts of <0.1 cm diameter. (<40 mil) per sample. These impacts did not affect tensile strength as much as the atomic oxygen erosion (see previous discussion), but the potential effect of many impacts on a composite structure is obvious. Ground simulationsin the NASA/Marshall Space Flight Center Space Debris Impact
815
Facility have been performed on a number of composite materials. These materials were Kevlar 49/epoxy, non-impregnated Kevlar 49, Spectra 900/epoxy and IM6/3501-6 graphite/epoxy. The test configuration consisted of a classic Whipple design with the composite material used as a bumper protecting a 2219-T87 aluminum pressure wall sample. The results of these tests (Schonberg, 1990) indicated that some composite systems, such as the Kevlar 49 cloth and Spectra 900/epoxy, performed better than equivalent weight aluminum systems as measured by damage to the pressure wall. However, the Kevlar 49/epoxy and IM6/3501-6 graphite/epoxy panels did not perform as well as an equivalent weight aluminum system at protecting the pressure wall plate from damage, but damage by spallation was far less for composite than metal bumper plates. The composite bumpers did experience large areas of delamination, though this was restricted to the outer layers with peeling in the direction of the surface laminate fibers. Some charring and melting did occur with the Kevlar 49 cloth (Fig. 36.4) and Spectra/epoxy test articles.
I;
‘I ..c
-.
Fig. 36.3 Cross-sectional photomicrograph of coated graphite-reinforced OMC showing impact damage and post-impact atomic oxygen erosion.
Fig. 36.4 Simulated meteoroid/debris impact in Kevlar 49 cloth panel.
816 Environmental efects on Composites Particulate radiation
Thermal extremes
Particulate or ionizing radiation in the space environment consists of trapped radiation in the Van Allen radiation belts, solar and galactic cosmic radiation and bremsstrahlung radiation. Protons, electrons, neutrons, alpha particles, X-rays, gamma rays and other charged particles bombard spacecraft in orbit. Bremsstrahlung radiation is ionizing radiation resulting from the deceleration of electrons. Various computer models exist for predicting the radiation dose during a mission at a particular orbit, but the high variability of solar activity and the resulting magnetic storms may give radiation calculations an uncertainty of an order of magnitude or more. Protons, electrons and cosmic rays may affect composite materials by surface chemical changes. The radiation threshold for observing effects is between 108 and 109 rads for graphite/epoxy and graphite/polysulfone. In a cooperative effort between Boeing and Langley Research Center (Fogdall, Russell and Cannaday 1980), these materials were irradiated with high-energy protons and electrons (approximately 1MeV) then tested for flexural durability. The composite systems tested were T300/934, T300/5208, and C6000/P1700. Radiation doses of 5 x lo9 rads or more resulted in surface blistering of the T300/934, although this had a limited effect on mechanical properties. Flexural testing of irradiated samples showed no significant changes for low radiation doses. A follow-on study performed at Boeing (Fogdall et al., 1983) involved high radiation doses to T300/934, C3000/934, C3000/PMR15 and C3000/P1700 composites. In this test, compressive breaking strength at a 45" fiber orientation and dynamic mechanical analysis (DMA)were used to measure the effect of 1O1O rads of electron radiation. As much as 15% degradation in breaking strength was noted. The DMA tests also showed a decrease in glass-transition temperature for irradiated samples.
As a spacecraft orbits, it moves from sunlight to shadow. Solar radiation, both direct and reflected from the earth (albedo), and thermal radiation from the earth cause temperature extremes of -185 to +150"C (-300 to +300"F). Spacecraft heat sources, such as engines, electronic equipment and batteries, also contribute to the thermal environment. Active thermal control systems or protective coatings with optimized solar absorptance and infrared emittance can minimize thermal extremes. Solar absorptance is defined as the ratio of absorbed light to incident light, and infrared emittance is defined as the ratio of emitted heat to input heat. Low absorptance values reduce the maximum temperatures when exposed to direct or albedo radiation and low values of emittance reduce minimum temperatures when exposed to deep space. Stability of these properties is important for long-term exposures. Thermal cycling effects on composites will be discussed later in this chapter. Spacecraft leaving and returning to earth must endure heating during ascent and reentry through the atmosphere. Re-entry temperatures can reach 1482°C (2700°F). Reinforced carbon-carbon composites are currently utilized in the Space Shuttle nose-cap cover and wing leading edges. Another application under investigation is a composite nose cone for the Space Shuttle External Tank (Sigur and Gray, 1990). This design not only weighs less than an equivalent metallic design but also can maintain the internal environment within desired temperature limits without thermal insulation. Temperatures on the Space Shuttle External Tank nose cap may range from -150°C (-297°F) around the liquid oxygen vent louvers to 500°C (930°F) during ascent. Mechanical property tests, including tensile, compression, in-plane shear and interlaminar shear tests, have been performed on laminates, moisture saturated per ASTM-D618 to reproduce the high humidity at the launch site. These composite materials were also
Environments and efects 817 tested for 'ply lift' or moisture-induced delam- Because moisture is removed from the comination at high temperatures. Several posite in a vacuum, mechanical properties, composite systems were studied, including such as compressive and interlaminar shear graphite/bismaleimide and graphite/PMR- strengths, may be improved through expo15, but Celion G-30/500-3K graphite/BASF sure. However, the designer must consider 506 phenolic was chosen because of its dura- possible dimensional changes due to moisture bility during the ply lift testing. desorption. Dimensional stability is critical for space hardware such as optical benches and truss structures. Another design concern is the Ultraviolet radiation effect of outgassed moisture on sensitive Ultraviolet radiation is that band of light from optics or electronic equipment present on a 300 to about 4000 A. Ultraviolet radiation may spacecraft. Vapor barriers of metallic foils have cause degradation through molecular weight been used to prevent line-of-sight deposition change and cross-linking in the resin system. of moisture and other outgassing products. However, this damage is generally limited to The UTIAS experiment flown on LDEF darkening of the resin in the surface layer. (Tennyson, 1991) contained a variety of flat Figure 36.5 is a photomicrograph of a LDEF and tubular composite samples consisting of composite laminate exposed only to UV. T300/5208, T300/SP-288 and T300/934 Coatings, such as thermal control tape, have graphite/epoxy samples, boron/SP-290 epoxy been used to protect composite materials from and Kevlar/SP-328 epoxy. A data acquisition degradation. system recorded outputs from 16 strain thermal gages attached to the composite samples for 371 days. Outgassing was measured, ranging from 40 days for the T300/934 to 120 days for the Kevlar/SP-328. Coefficient of thermal expansion increased slightly for the 0" configuration (Table 36.1). This change should be considered when designing zero CTE laminates for space applications. 36.3.6 TEMPERATURE
Temperature effects on composite materials discussed in this section include cryogenic temperatures, elevated temperatures and thermal cycling between these extremes. Cryogenic temperatures do not appear to Fig. 36.5 Cross-sectional photomicrograph of UV- affect the mechanical properties of exposed LDEF graphite-reinforced OMC showing graphite/epoxies or graphite/polyimides sigminimal degradation. nificantly. Elevated temperatures for a prolonged Vacuum and outgassing period of time can seriously affect the properOrbital atmospheric pressure varies according ties of a composite, with even greater effect if to altitude and solar activity. Average pressure moisture is present. Susceptibility to matrix is generally 133 x N/m2 (1 x torr) in softening is not only dependent on the resin N/m2 but also the lay-up. A study of graphite/polylow earth orbit, decreasing to 133x (1 x lO-I4torr) at geosynchronous orbit. imide properties used two different lay-ups, a
818 Environmental effects on composites Table 36.1 Summary of LDEF/UTIAS composite material thermal data (Tennyson, 1991)
Strain gage number
Thermal &?age number
Ambient" CTE (1PPF)
Space CTE (1P P F )
Thermal vacuum facility' CTE (lOd/W
1
1
9.84
10.0
10.0
Graphite/Epoxy T300/934
2 (0") 3 (90")
2
1.32 14.7
3.33 13.9-15
2.50 16.1
Kevlar/Epoxy 93-328
4 (90") 5 (0°)
3
33.9 0.10
30-35 0.71
35.3 0.63
Graphite/Epoxy T300/SP388
6 (90") 7 (0°)
4
14.6 0.97-1.57
13.6-14.3 -1.14-3.33
15.4 3.75
Graphite/Epoxy T300/5208
8 (90")
5
15.6
12.5-15.3
16.1
Boron / Epoxy SP-290
9 (GO") 10 ( 6 0 )
6
1.57 11.7
1.67-2.22 7.5-11.1
0.44-2.0 12.7
Material type
Stainless Steel (calibration)
Configuration
At atmospheric pressure prior to launch. Measured in space environment on LDEF during first 371 days in orbit. Measured in laboratory thermal vacuum test facility (22 h at 133 x l t 5N/mZ (1 x l t 5torr), -40°F to 150°F) after 2114 days in orbit and 184 days at ambient conditions. a
0" unidirectional lay-up and a (0, +45,90), lay- composites, and advanced carbodcarbon up, tested at temperatures ranging from -157 composites. to +315"C (-250 to +600"F) (Lisagor, 1979).Test Temperature effects are not limited to the results showed little change in interlaminar matrix material. Extended operation at 350°C shear strength for the quasi-isotropic lay-up (660°F) and 450°C (840°F) can cause oxidation while the 0" unidirectional samples dropped of low-modulus PAN-based fibers and high to approximately 40% of original strength at modulus PAN- or pitch-based fibers, respecelevated temperatures. Reduction of normal tively. Oxidation resistance can be improved moisture content by vacuum drying reduced with higher purity fibers. the loss to only 70% of room temperature Thermal cycling conditions are common for strength. a number of applications, including aircraft High-temperature resins under develop- and spacecraft. Thermal cycling may induce ment, such as AFR700B developed by the Air microcracking in some composites. A study of Force (Brown, 1991),have reached glass transi- this microcracking behavior in graphite/PMRtion temperatures of 416°C (780°F). AFR700B 15 composite materials was performed at Rohr retains 50% of its mechanical properties up to Industries (Sullivan and Ghaffarian, 1988). 370°C (700°F). PMR-15 is another high-tem- Woven laminates of C3000/PMR-15 and unidiperature resin with excellent properties, but rectional tape lay-up of C6OOO/PMR-15 were large or thick structures require debulking a thermally cycled between -18 and +232"C (0 few layers at a time due to the high volatile and 450'F) up to 2000 cycles. The C3000/PMRcontent. Other materials being developed for 15 developed microcracks, with the number of high-temperature applications include tita- microcracks dependent on the number of thernium matrix composites, ceramic matrix mal cycles. Decreases in compressive and shear
Protective coatings 819 strengths were noted, although tensile strength, a fiber-dominated property, was not significantly affected. The unidirectional tape lay-up, C6OOO/PMR-15 did not crack during testing. 36.4 PROTECTIVE COATINGS
When an environmentally resistant composite material cannot be utilized, protection of the material through the use of coatings is necessary. A variety of coatings have been developed for protecting composites from various environments. Standard marine paint, pigmented gel coatings and polyurethanes have been used to prevent ultraviolet damage and weathering erosion of marine composites. The following coating examples have been used in space but
may be applicable to other environments. On the LDEF, S-901 glass/epoxy composites were exposed to the space environment. Three of these samples were flown with thermal control tape, consisting of 50.8 pm (2 mil) aluminum with 50.8 pm (2 mil) pressure-sensitive SR574 silicone adhesive. The glass/epoxy samples without tape suffered more mass loss than the protected samples, and solar absorptance of the unprotected samples increased 5.4%, probably due to ultraviolet radiation. Post-flight peel tests of the thermal control tape showed an increase in strength with some embrittlement of the adhesive. Figure 36.6, a SEM photograph, also shows the degree of micrometeoroid/debris protection provided by the tape.
. .- b
Fig. 36.6 Micrometeoroid/debris impact in thermal control tape. Substrate is S-glass/epoxy.
820 Environmental effects on composites A 60nm (600A) Si0,/100nm ( 1 O O O A ) nickel vapor-deposited coating appears to have prevented resin loss in graphite/epoxy samples flown on LDEF (Young et al., 1991). Inspection of selected composites (C6000/P1700, C3000/P1700, T300/934 (with two different fiber areal weights) and T300 /5208) revealed dramatic visual effects, material loss and a deterioration in mechanical performance for unprotected composites. Chemical characterization suggests that there is no significant change in molecular structure of the surviving polymeric matrix resin in these composites. No change in glass transition temperature was apparent. The majority of the unprotected flight samples showed a decrease in tensile strength and modulus. Coating/composite systems for large space truss structure applications have been tested in ground simulations of atomic oxygen, ultraviolet radiation, and thermal cycling (Dursch and Hendricks, 1987; Dursch and George, 1993). Coatings studied were chromic and phosphoric acid anodized aluminum foil, sputtered SiO,/sputtered Al/Al foil, bare A1 foil, electroplated nickel with and without SiOxcoatings and inorganic sol gel solutions. P75S/934 composite was used to form the truss tubes. The aluminum foil used was 1145-H19 aluminum alloy with a thickness of 50.8 pn (2 mil). Thicker foils can be used, depending on weight constraints and flexibility of any coatings applied to the foil. The foil was bonded to the graphite/epoxy truss tube with an epoxy film adhesive and was co-cured. The chromic acid anodized foil provided good protection as well as optical properties. The phosphoric acid anodized foil also provided good protection while maintaining adhesion. Desired optical properties were obtained by varying thickness of silicon dioxide and sputtered aluminum on top of the aluminum foil. The optimized thicknesses were 1pn (0.04 mil) SiO, and 300 nm (3000 A) Al. The electroplating process was chosen for study because of its low application costs,
good corrosion resistance and good uniformity. Irregular shapes, such as end fittings, may also be coated with this process. The composite surfaces were sanded prior to plating to improve adhesion. However, atomic oxygen attack resulted in loss of adhesion in the nickel coating without SiOxovercoat. Numerous other coatings have been studied for space environment resistance. Braided and double-braided aluminum coverings, indium-tin eutectic coatings, various siliconebased paints and zinc oxide pigment in a potassium silicate and silicone elastomer binder have been tested (Piellisch, 1991). A proposed design for solar array backing on the NASA space station employs woven E-glass sandwiched between two layers of silicon dioxide-coated Kapton.
REFERENCES
Anon. 1970.Meteoroid Damage Assessment. NASA SP-8042 Ashbee, Ken. 1989. Fundamental Principles of Fiber Reinforced Composites. Lancaster, PA: Technomic Publishing. Beckmann, Hans-Dieter and Oetting, Hermann. 1985. Fiber reinforced plastics move inside engine. Automotive Engineeying 93(5):34-41. Brown, Alan S. 1991. The Air Force finds an ultrahigh temperature resin. Aerospace America 29 (10):56-57. Dexter, H. Benson. 1987. Long Term Environmental Effects and Flight Service Evaluation of Composite Materials. NASA TM-89067. Dobyns, Alan. 1991. Structures. Aerospace America 29(12):38-39. Dursch, Harry W. and Hendricks, Carl L. 1987. Protective coatings for composite tubes in space applications. S A M P E Quarterly 19(1):14-18. Dursch, Harry W. and George, P. 1993. Composite Protective Coatings for Space Applications. Paper read at Third LDEF Post-Retrieval Symposium, November 8-12, 1993, at Williamsburg,VA. Fogdall, L.B., Russell, D.A. and Cannaday, S.S. 1980. Radiation Exposure of Composite and Polymer Materials. Report for Contract NAS1-15606. Fogdall, L.B., Lindenmeyer, P.H., Sheppard, C.H., Russell, D.A. and Cannaday, S.S.. 1983.
References 821 Development of Facilities, Quality Control Procedures, and Testing Techniques for Irradiation of Spacecraft Composite Materials. Report for Contract NAS1-16854. Fried, N. 1969. Marine Applications. In Handbook of Fiberglass and Advanced Plastics Composites, (ed. George Lubin) pp. 747-783. New York: Van Nostrand Reinhold. Graner, William R. 1982. Marine Applications. In Handbook of Composites, (ed. George Lubin) pp. 514-529. New York: Van Nostrand Reinhold. Kamenetzky, R.R. and Whitaker, A.F. 1992. Performance of Thermal Control Tape in the Protection of Composite Materials to Space Environmental Exposure. NASA TM-103582. Kessler, D.J., Reynolds, R.C. and Anz-Meador, P.D. 1989. Orbital Debris Environment for Spacecraft Designed to Operate in Low Earth Orbit. NASA TM-100471. Levine, Arlene, (ed.). 1991. LDEF-69 Months in Space: First Post-Retrieval Symposium. NASA CP-3134. Levine, Arlene, (ed.). 1992. LDEF-69 Months in Space: Second Post-Retrieval Symposium. NASA CP-3194. Lisagor, W. Barry. 1979. Mechanical Property Degradation of Graphite/Polyimide Composites After Exposure to Moisture or Shuttle Orbiter Fluids. In Graphite/Polyimide Composites, (ed. H. Benson Dexter and John G. Davis, Jr.) pp.273-287. NASA CP-2079. Piellisch, Richard. 1991. New solar arrays mean new materials. Aerospace America 29(5):20-23. Pilpel, Edward D. 1982. Expanded Design Analysis of the Use of Composites in Determining Snow Ski Characteristics. In Materials Overview for 2982, SOC.Advanc. Mater. Proc. Engng, pp. 616-627. Schonberg, W. 1990. Hypervelocity Impact Testing of Non-metallic Materials. 17th Cong. Intern. Council of the Aeronaut. Sci., Sept. 9-14, 1990, at Stockholm, Sweden, paper ICAS-90-1.7.3. Sigur, W.A. and Gray, C.R. 1990. Composite Applications on the External Tank. Final Report from Martin Marietta Manned Space Systems, NASA Contract NAS8-33708.
Springer, George S., Sanders, Barbara A. and Tung, Randy W. 1981. Environmental Effects on Glass Fiber Reinforced Polyester and Vinylester Composites. In Environmental Effects on Composites Materials, pp. 126-144. Westport, CT Technomic Publishing. Spry, William J. 1987. Sports and Recreational Equipment. In Engineered Materials Handbook, Vol. 1, (ed. Theodore J. Reinhart et al.) pp. 845-847. Metals Park, Ohio: ASM International. Staunton, R. 1982. Environmental Effects on Composites. In Handbook of Composites, (ed. George Lubin) pp. 514-529. New York: Van Nostrand Reinhold. Sullivan, Lawrence J. and Ghaffarian, Reza. 1988. Microcracking Behavior of Thermally Cycled High Temperature Laminates. Paper read at 33th Intern. SAMPE Symp. Exhib., March 7-10, 1988, Anaheim, CA. Tennyson, R.C. 1991. Composite materials in space - results from the LDEF satellite. Canadian Aeronautics and Space J. 37(3):120-1 33. Vette, J.I., Lucero, A.B., Wright, J.A., King, J.H. and Lavine, J.P. 1967. Models of the Trapped Radiation Environment. NASA SP-3024. Whitaker, A.F. 1991. Coatings Could Protect Composites from Hostile Space Environment. Advanced Materials and Processes 139(4). Whitaker, A.F. and Young, L.E. 1991. An Overview of the First Results on the Solar Array Materials Passive LDEF Experiment (SAMPLE), A0171. Paper read at First LDEF Post-Retrieval Symposium, June 2-8,1991, Orlando, FL. Whitaker, A.F. 1991. Preliminary Assessment of LEO Effects on LDEF Experiment A0171 Composite Material Surfaces. Paper read at LDEF Materials Workshop, Nov. 19-22,1991, at Langley VA. Young, Philip R., Slemp, Wayne S., Witte, Jr., William G. and Shen, James Y. 1991. Characterization of Selected LDEF Polymer Matrix Resin Composite Materials. Paper read at 36th International SAMPE Symposium and Exhibition, April 15-18,1991, at San Diego, CA.
SAFETY AND HEALTH ISSUES
37
Jennifer A. Heth
37.1 INTRODUCTION
On a weekly basis, there are conflicting media reports indicating that something is good or bad for us, depending on the study done, the topic and the amount of press given to it. However, by reviewing the information and adding common sense and moderation to our lifestyles, we are able to discern good from bad, thus allowing us to lead healthy and safe lives. Working with composite materials, or any chemicals for that matter, should be viewed in the same manner. Know what material is used, how it is handled, what is known about it and how to reduce the risk of injury from any hazard associated with it. From a safety perspective, proven engineering and administrative techniques and controls exist that can make the workplace safe if implemented correctly. There are fairly universal industrial hygiene protocols that merely need to be implemented to effectively minimize potential exposure to any hazard, whether it be chemical or physical. On health issues, there are numerous toxicological papers published on chemicals used in the composites industry. However, without a toxicological background or proper analysis and interpretation of the data, it is difficult to know what the studies’ conclusions are and, more importantly, if they are valid. To assist users in understanding safety and health issues for composite materials, the
Handbook of Composites.Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Safety and Health Subcommittee of the Suppliers of Advanced Composite Materials Association (SACMA) researched and published Safe Handling of Composite Materials (3rd Edn, April 1996). This chapter is an edited version of the booklet, with new information added as appropriate. The information is not meant to be inclusive. Rather, the reader should be aware of the issues addressed and make further investigation as needed. Note that data cited was current in 1997. 37.2 HEALTH INFORMATION TERMINOLOGY
It is important to understand the terminology in order to read and assess toxicological data. This section will concentrate on some of the basic terms and definitions that are applicable to the composite user, especially in reading Material Safety Data Sheets, the most common source of information on materials. 37.2.1 TOXIC, HAZARD AND RISK
These three words are used frequently and sometimes interchangeably. To avoid confusion, they are defined as follows.
Toxic refers to a poison or poisonous substance that may cause a harmful effect in the body. A substance’s toxicity characteristic is a property of the chemical, similar to its color or odor. This is as true for chemicals like water, salt or sugar as it is for cyanide, snake venom or botulism toxin (Fig. 37.1).
Health information terminology 823 ‘All chemicals are toxic, there
\
are none which is not ...
/
The right dose differentiatesa poison from a remedy.’ /
,6th Paracelsus Century
Risk describes the probability or likellhood that a hazard will result in a harmful effect. Regardless of the toxicity, hazard or risk associated with a substance, care should always be taken to minimize exposure. This is because what is known today may be inadequate compared to what tomorrow’s science may discover (e.g. asbestos usage and asbestosis or tobacco and lung cancer). 37.2.2 ACUTE TOXICITY VERSUS CHRONIC TOXICITY
Fig. 37.1 Toxicology principle.
When discussing toxicity, reference will typically be made to two types: acute or chronic. One can be instantaneous and the other manifests itself over a long period of time.
Hazard takes into account the toxicity of a substance along with exposure to it. Something can be extremely toxic, but if there is no exposure, there is no hazard (Fig. 37.2). Conversely, it is important not to be negligent in handling low toxicity substances, as extreme exposure (high concentration or long exposure time) could result in high hazard.
Acute toxicity occurs when a harmful effect is experienced after a single or short-term exposure to a substance. It usually occurs instantly or within a short time period. Typically, exposure will be by skin or eye contact or inhalation. In industrial settings, it would be unlikely for exposure to be by ingestion, though this is another route.
=
K
Minimal hazard
Low exposure
High toxicity
AM 6
0 9-
X Low toxicity
Fig. 37.2 Risk assessment equations.
nw
Extreme exposure
= Elevated hazard
824 Safety and health issues The most common measure of acute toxicity (possibly years or decades). Because chronic through the oral or dermal routes is called the effects develop slowly and are measured later, median lethal dose or LD,,. The LD,, (mg of and it is not possible to keep the workplace chemical per kg of body weight) is the amount toxic-free, it is important to eliminate or miniof material that kills 50% of a group of experi- mize exposure. mental animals, usually mice or rats (Fig. 37.3). The measure by inhalation is LC,,, or 37.2.3 SENSITIZATION median lethal concentration. It is expressed as an airborne concentration in milligrams of In some cases, an allergic reaction to a subchemical per cubic meter (mg/m3) of air, or stance will develop with one exposure, or over time with repeated exposures. This is called parts per million (ppm) in the air (Fig. 37.3). sensitization. Once a person is sensitized to a substance, the extent of the reaction does not 0 Oral and dermal routes: necessarily relate to the degree of exposure. Median lethal dose - LD, (mg/kg) Also, people who are sensitized to one chemiA large LD, (e.g., 5000 mglkg) equals a low degree cal may react to other similar materials. This is of acute toxicity, typically relating to a low health known as cross sensitization (Fig. 37.4). hazard. Certain individuals may be sensitized to some chemicals used in the composites industry. For 0 Inhalation route: example, there have been reported cases of epoxy sensitization. In these instances, it is Median lethal concentration - LC,, (rng/rn3) important that an employee not be exposed For rats, a four hour exposure period is commonly further, or allowed the opportunity to be used. Because other time durations are used, data exposed. To do this, engineering or adminiswill usually be reported with time specified. trative controls should be implemented. If these are not possible or practical, then the Fig. 37.3 Acute toxicity measures. employee should be removed from chemicals causing reaction. The acute inhalation hazard is dependent on the material's toxicity, together with its physical properties, such as vapor pressure for gases 37.2.4 EXPOSURE LIMITS or particle size for aerosols, particulates and There are limits or values (Fig. 37.5) estabdusts. Knowledge of the physical properties lished by health groups to assist the user in will assist in determining whether a material is controlling exposure to certain chemicals; likely to become airborne, thus inhalable. some are regulated (OSHA's Permissible Chronic toxicity occurs when adverse health Exposure Limit, PEL), others are recomeffects are manifested after exposure to a sub- mended (ACGIH's Threshold Limit Value, stance over a long period of time (e.g. TLV). These limits are based on actual industhroughout a working lifetime) or by long-term trial experience, experimental animal and effects resulting from one or a few doses. These human studies and, when possible, a combieffects can occur following repeated exposures nation of all three. Exposure limits are developed for protecto chemical substances through dermal or ocution against serious health effects or irritation, lar contact, halation or ingestion routes. Chronic toxicity testing includes systemic narcosis, or nuisance. They are intended for toxicity, mutagenicity carcinogenicity, repro- the control of potential health hazards in the ductive toxicity and sometimes epidemiological workplace. Because each individual is differstudies, all of which are very time consuming ent, there will be incidents of people affected
Industrial hygiene 825 Once sensitized (have an allergic reaction), the following may occur:
fl 0
9.
X Low exposure
0 .
Low toxicity
fl 0 9.
X
0 .
Low toxicity
Chemicals similar to substance causing sensitization
nw
Extreme exposure
= Allergic reaction
= Allergic reaction
X Low exposure
Fig. 37.4 Sensitization equations.
by some substances at concentrations below 37.3 INDUSTRIAL HYGIENE the PELs or TLVs (Fig. 37.5). This is because the Potential hazards can be controlled through the limits/values are based on the exposed popu- implementation of sound industrial hygiene lation being 'normal', and does not address practices. Whether employees are working in a aggravation from pre-existing conditions, ill- high or low hazard area, the recognition, evalunesses or lifestyle choices. It is best to maintain ation and control of conditions are necessary to concentrations of atmospheric contaminants at avoid occupational illness or discomfort. This is the lowest level reasonably possible through industrial hygiene. process or engineering controls. It is important to apply good engineering, When referencing PELs and TLVs, a 'skin' process and administrative controls in the worknotation may be assigned. This indicates that place along with effective industrial hygiene dermal exposure should also be considered practices (e.g. monitoring, surveys, proper perwhen evaluating exposure to that substance. sonal protective equipment usage). Without Always check the current OSHA standards for addressing the fundamentals, it is difficult to PELs and the latest edition of the ACGIHs maintain employee comfort and, hence, protecTLVs: Threshold Limit Values and Biological tion against exposures that may be present. Exposure Indices, along with the supplier's Industrial hygiene practices for composites Material Safety Data Sheet, for information on are common to those for other industries. limits and protection against exposure.
826 Safety and health issues Exposure limits PEL
(Permissible Exposure Limits) Airborne exposure limits issued by the OccupationalSafety and Health Administration (OSHA), 29 CFR 1910.1000, which are legally binding.
TLV
(Threshold Limit Values) Recommended exposure limits for airborne concentrations of substances. They are issued by the American Conference of Governmental Industrial Hygienists (ACGIH) and published annually.
Exposure limit descriptions Both PEL and TLV are expressed in similar terms describing exposures: TWA
(Time Weighted Average) This represents a normal 8 h day and a 40 h work week, to which nearly all workers may be exposed without adverse health effects. In literature, data will appear as PEL-TWA or TLV-TWA.
STEL
(Short-Term Exposure Limit) Concentrations to which people can be exposed for a short period of time without suffering from 0 0
irritation; chronic or irreversibletissue damage; or narcosis of sufficient degree to increase likelihood of accidental injury, impairment of self-rescue or reduction of work efficiency
provided the daily PEL or TLV-TWA is not exceeded. Data will appear as PEL-STEL or TLV-STEL in written works. C
(Ceiling) The concentration not to be exceeded during any part of the working exposure. This will appear as PEL-C or TLV-C in documentation.
Fig. 37.5 Exposure limit terminology.
Periodic exposure assessments should be conducted taking into consideration routes of exposure associated with composite use (Table 37.1). Monitoring should be routine, and on demand whenever unusual odors are noticed, visible contamination is heavier than normal,
or a new process is to be used. After any new process or modification of an old one, good work practice is to ensure that existing controls to prevent exposure are implemented and effective. Two potential exposures posed by composite
Table 37.1 Routes of exposure for composite users Skin and eyes
Typically hands, lower arms and face are exposed. However, if personal hygiene is not good, other areas of the body may be susceptible to exposure. Avoid exposure, especially in cases where dermatitis or sensitization has been confirmed. Avoid contact with chemicals that can be absorbed through the skin, as systemic and/or local effects may occur.
Inhalation
Good ventilation will minimize possible exposure from the release of solvents or dusts generated.
Ingestion
Thorough washing of the hands prior to eating or smoking provides sigmficant protection from the effects of accidental ingestion.
Injection
Needles and sharps are not normally a concern. However, shards from cured composites or brittle fibers, or needles from weaving or sewing processes can puncture the skin and chemicals could enter the body.
Industrial hygiene 827 usage are skin contact with materials that could result in irritation leading to dermatitis or sensitization, and inhalation of particulates from operations such as cutting, grinding and finishing. Both concerns can be eliminated or minimized with implementation of proper gloves/clothing, good ventilation and process conditions, and effective training. 37.3.1 ADMINISTRATIVE CONTROLS
In the National Safety Council’s Fundamentals of Industrial Hygiene (3rd Edn), administrative
controls are defined as ’methods of controlling employee exposures by job rotation, work assignment, or time periods away from the hazard’. These controls may be implemented when e n p e e r i n g controls cannot reduce the exposure to permissible levels (Fig. 37.6). Administrative controls also encompass ’other work rules’, such as company policies (Table 37.2). One example of a process hazard in the composites industry, controllable through engineering and administrative rules, is an out-of-control exothermic reaction. This is an unintentional runaway chemical reaction of a
Table 37.2 Workplace tools to prevent exposures 1. Safety controls
Confined space entry; hot work permits; lockout/tagout procedures; pipeline breaking permits; safe work permits.
2. Personal controls
Knowledge and use of Material Safety Data Sheets (MSDS);company policies and procedures; regulation of employee’s exposure time in a work area.
3. Education and training
Training in the use and handling of materials and any associated hazards; understanding of the training.
4. Materials handling Information from MSDS on proper handling of materials
5. Process controls
Isolation of process controlled work areas; eating, drinking, smoking should be separate from any work area as should food storage; control of process-related hazards, such as an out-of-control exothermic reaction potential. There should be procedures, equipment, ventilation systems and process hazard analysis and training in place; wet processes to reduce dust generation.
6. Personal hygiene
Thorough washing of hands before each job break and eating/drinking/smoking and toilet use along with use of good hand creams after each wash.
7. Warnings and labels
Proper labels on chemicals; follow OSHA’s Hazard Communication Standard.
8. Housekeeping
Keep work areas clean and free of clutter; remove dust by vacuuming instead of blowing it off work areas into the air.
9. Emergency plans
Plans and procedures for dealing with any chemical emergency; personnel should have the knowledge, skills and equipment to respond appropriately.
10. Ventilation
General, diluted (supplied air) and local exhaust ventilation.
11. Medical controls
Good occupational medical program and protocol.
12. Engineering
Design, maintenance and hazard studies; shielding, monitoring devices, and alarm systems.
practices
13. Personal protective Respirators, gloves, body suits, boots, jackets, goggles, eyewear, face shields,
equipment 14. Administrative
controls
safety shoes/boots, headgear. Job rotation, scheduling machine times to reduce number of workers exposed, scheduling work times to avoid exposures (eg. cooler times of day to avoid heat stress), reduction of work periods.
828 Safety and health issues
I
1
Exothermic reactions can be avoided and minimized by clearly defining and following the proper procedures, thoroughly training employees who work with resins and prepreg materials, making certain that equipment is in good working order, and that safety devices or
Personal protective equipment (PPE)
benzyl phthalate, ABP) are in place and func-
Engineering controls
I
J.
-Administrativecontrols
I 1
Fig. 37.6 Industrial hygiene order of priorities.
resin system, alone or in prepreg form, typically called an exotherm. It may occur under any of these conditions: 0 0 0 0
heating or mixing a resin too long; heating a resin too fast; allowing a resin to get too hot; contamination or mislabeled chemicals.
These factors have been known to start or extend an exotherm: 0 0 0 0 0 0 0 0 0 0
resin mass scale up and equipment cannot dissipate heat; deviating from procedures; disabling safety equipment; casting hot melt resins too deep; process equipment malfunction; variability in raw materials; mixing incompatible chemicals and curing agents; contaminating chemicals, e.g. poor housekeeping; uneven dispersing or mixing of chemicals; trying to mass cure resin oE prepreg in an oven or autoclave.
37.3.2 PERSONAL PROTECTIVE EQUIPMENT (PPE)
Personal protective equipment is used to control exposures when they cannot be reduced to an acceptable or practical level using engineering and administrative controls. In the composite industry, there are three main exposure areas that may require PPE to protect the skin, eyes and lungs. 37.3.3 SKIN PROTECTION
Hand contact exposure is most common when using composites. The use of proper protective gloves is important to control exposure. Along with gloves, other equipment such as jackets, arm protectors, or body suits may be necessary depending on the process or potential for exposure. Gloves Glove selection should take into account the parameters in Table 37.3 to ensure the correct
Table 37.3 Proper glove selection criteria Chemical resistance
Glove must be impermeable to the chemical being used. An incorrect glove choice may allow the material to come in contact with the skin.
Dexterity and comfort
If the user cannot work comfortably with the glove, it will not be used.
Glove lining
Lining material may cause irritation or excess sweating. Consider liners or other alternatives.
Glove surface
The outer surface may contain residual coatings that may contaminate and possibly delaminate a composite part.
lndustrial hygiene glove choice and use by the employee. Additionally, Table 37.4 gives a brief listing of glove types used in the composite industry and what they protect against. SACMA's companion video to the Save Your Skin! booklet demonstrates correct glove selection and usage for composite users. It is important to know that there is not an 'all purpose' glove for protection from all exposure. Proper protection may require wearing one type of glove over another. Basically, there are two glove types:
829
aramid protect well against heat, cuts or scrapes. Glove choice must depend on the job and its duration.
Skin creams
There are creams that are applied directly to the hands and forearms for protection. Two types of cream discussed in Table 37.5 indicate the difference between moisturizers (no protection against chemicals) and barrier creams (limited protection). Barrier creams may be used alone or in combination with gloves. 1. Chemically resistant: When using barrier creams, there is a possiA selection of gloves used for protection against exposure to chemicals. Depending bility of contamination to the composite on the permeability of the glove material to material. If there is a heavy reliance on the use the chemical used, the glove chosen will of barrier creams, process engineers need to protect the wearer for only a limited time. investigate alternatives in engineering, process Even the proper glove does not last forever. changes or alternate PPE. Since barrier creams are not moisturizing creams, a good moisturiz2. Mechanically or thermally resistant: These gloves, such as leather, cotton or ing cream should also be applied after work to
Table 37.4 Examples of glove types Glove type
Hazard ...
Aramid
Cutting, intense heat
Cotton
Abrasions
Disposable plastic (Latex)
Microorganisms, mild irritants, fibers
Natural rubber (Latex)
Acetone (< 1 h), epoxies, methyl ethyl ketone (MEK), light work
Leather
Abrasions, punctures, fibers
Metal mesh
Cuts, scrapes
Neoprene
Acetone (< 1 h), solid epoxies, DMSO, isopropyl alcohol
Nitrile
Solid epoxies, TDI, isopropyl alcohol
Polyvinyl alcohol
Methylene chloride, bisphenol A epoxies, toluene, MIBK, TDI, styrene, THF (< 4 h), l,l,l-trichloroethane
Polyvinyl chloride
TDI, isopropyl alcohol, (< 4 h)
Rubber, butyl
Isopropyl alcohol, Dimethylformamide (DMF), DMSO. MEK, MIBK, acetone, N-methyl-2-pyrrolidone (NMP)
Rubber, insulated
Electrical shocks and burns
VITON@
Toluene, xylene, styrene, isopropyl alcohol, l,l,l-trichloroethane
Sources: Save Your Skin!, SACMA 1990; Safe Handling of Advanced Composite Materials, SACMA 1996.
830 Safety and health issues Table 37.5 Skin creams
Moisturizing creams
Replenish moisture in the skin which is lost after washing up. Use regularly to avoid dry skin conditions which could lead to dermatitis. Gloves must always be worn when moisturizing creams are used.
Barrier creams
1. Water repellent creams: Leave a thin film of lanolin, beeswax, petroleum or silicone on skin. Used mainly in machine shop operations. 2. Solvent repellent creams: Leave a thin film on the skin, visible or not, which will repel oils, paints, and solvents. Barrier creams do wear off and must be
reapplied for constant protection.
keep skin healthy. Additionally, barrier creams may actually increase abrasion from fiber dust, so their effectiveness needs to be monitored.
37.4 POTENTIAL HEALTH AND SAFETY HAZARDS IN COMPOSITE PROCESSES
When designing equipment, processes and modifications, complete containment of vapors and dusts should be a goal. General 37.3.4 EYE PROTECTION ventilation should be provided to all work Eye protection should be selected based on areas, with local exhaust equipment designed impact (flying particles) and/or chemical to pull contaminants away from the splash possibilities. Selections should include employee's breathing zone. In almost every appropriate safety glasses, goggles, face composite process, the engineering emphasis shields or a combination of these. is on good ventilation to control solvent and dust exposures, along with other contaminants. 37.3.5 RESPIRATORY PROTECTION Figure 37.7 gives a summary of potential During operations such as resin mixing, health and safety exposures that could occur prepreg lay-up, machining or clean-up, respi- in composite processes. However, with good ratory protection may be necessary to reduce engineering, administrative and industrial exposures to vapors or dusts. hygiene practices, personal exposure can be There are two main types of respirators: air minimized. purifying and air supplying. Before using any Process improvements are driven by techrespirator, familiarity with OSHA's Respiratory nology and regulation (e.g. CA's South Coast Protection Standard (29CFR 1910.134) is essen- Air Quality Management District Rules 1171 tial. Not everyone can or should use a and 1128 are for emission reduction of volatile respirator. An employee must be medically organic compounds in coating and cleaning approved, fit tested, and trained to assure that operations). In the composites industry, some the respirator used is both appropriate and examples of minimized solvent vapors and protective. exposures include closed loop systems, cabin If using a cartridge (filter) respirator, the surroundings for processes, improved ventilacartridges are specific to hazards and must be tion designs and capture efficiencies, and maintained and changed periodically to pre- elimination of solvents or substitution of low vent exposure. If they are the wrong type, are vapor pressure solvents for high vapor presused too long, or become dirty, respirators are sure solvents. ineffective.
Toxicological properties of composite components 831 COMPOSITE PROCESSES
Fig. 37.7 Potential health and safety exposures in composite processes. 37.5 TOXICOLOGICALPROPERTIES OF COMPOSITE COMPONENTS
effects, the resin system is of primary concern. The reinforcement is secondary due to its lower hazard potential compared to resin systems. The supplier’s Material Safety Data Sheet Solvents are addressed as a separate category as (MSDS) and other sources of toxicological they are major sources for potential skin and information should be used for practical hazinhalation exposures. Some are part of the resin ard assessment on any material. Rarely will systems, while others are only used for cleandata be available on the composite system up purposes. itself because the methodology has not been fully developed to adequately test these kinds of materials. Therefore, the hazards of com37.5.1 RESINS posites are generally expressed in terms of the components’ hazards. This is a conservative Composites are by definition based on a resin approach that provides the user with the most system matrix applied to some reinforcing protective information because it is based on material. The resin system is often quite comthe most hazardous component which may plex, consisting of a basic resin (e.g. epoxy), only be a percent or less of the entire mixture. which is formulated with other materials, Typically, when discussing toxicological such as curing agents, diluents, accelerators,
832 Safety and health issues pigments or solvents. Components of the resin system may be supplied individually and formulated by the user, supplied as ‘Part A and Part B’ and blended before use, or supplied mixed as a ’one-pack’ system.
any product containing additives with known or potential health effects. Reference to the MSDS will advise you of hazardous materials. Some commonly used additives are listed in Table 37.7.
Epoxy resins
Other resin types
Epoxy resins, also known as glycidyl com- Over the years, epoxies have been the backpounds, are commonly used in composite bone of the composite industry. However, matrix resin systems. In addition to the stan- other resin systems play an important role and dard bisphenol A based epoxies, other are used for particular applications and perforglycidyl ethers, glycidyl esters, glycidyl mance properties. Table 37.8 discusses health amines, and epoxy novolacs are used in com- effects of some of these other resins. The associated health hazards may appear severe, but posite matrices. The greatest concern with epoxies is their exposures are avoidable through engineering potential to cause skin irritation, dermatitis or and industrial hygiene practices, especially skin sensitization, depending on the base good ventilation for control of vapors. epoxy used and the sensitivity of the individual exposed. Table 37.6 discusses epoxy based 37.5.2 REINFORCING MATERIALS resins and associated hazards. Most reinforcement materials (Table 37.9) are fibers, such as graphite or carbon, glass, Hardenerskuring agentskatalysts aramid or ceramic. Others may be used, but One or more of these additives are used to the application is typically specialized, resultenhance properties of composite materials. ing in small specialty runs. Overall, data Their percentage of the resin system may be indicates that most fibers have a low hazard small, but precautions ought to be taken for potential in initial form. However, they may Table 37.6 Health effects of various epoxies
EPOXY type
CAS nurnber(s)
Known health efects
Key notes
Bisphenol A based
1675-54-3, 25036-25-3, 25068-38-6, 25098-99-8
Possible skin sensitizer; low order of acute toxicity; slightly to moderately irritating.
Insufficient evidence to classify as a carcinogen according to IARC. Considering the many studies as a whole, the evidence does not show the resins to be carcinogenic.
Glycidyl amines
28768-32-3
Possible skin sensitizer; low order of acute toxicity.
Mutagenicity tests gave both negative and positive results.
Cycloaliphatics
2336-87-0, 30583-72-3
Irritant to skin and mucous membranes.
Not considered mutagenic or carcinogenic.
Glycidyl ethers
2210-79-9, 2426-08-6, 3101-60-8, 17557-23-2, 26447-14-3
Possible skin sensitizer. Moderate to severe skin and mucous membrane irritant.
Neopentylglycol diglycidyl ether has caused skin tumors when applied repeatedly to skin of shaved mice.
Toxicological properties of composite components 833
834 Safety and health issues
(.j
Toxicological properties of composite components 835 Table 37.9 Health effects of fibers
Reinforcement type
Known health effects
Key notes
Carbon or graphite fibers
Mechanical abrasion and irritation of the skin; possible dermatitis; physico-mechanical properties of the fibers rather than a toxico-chemical reaction. Possible reaction from the fibersizing. See resin health hazards.
PEL-TWA is 15 mg/m3 total dust, and PELTWA of 5 mg/m3 for synthetic graphite respirable dust. ACGIH has a TLV-TWA of 2 mg/m3 respirable dust for all forms of graphite except fibers. There are no limits for carbon fiber, though the US. Navy has set 3 carbon fibers/cc. EPA did not classify the potential carcinogenic properties of carbon fibers due to insufficient data.
Glass fibers
Mechanical irritation of eyes, nose, throat; possible skin sensitization, either from sizing or fiber.
Continuous fiber (used in composites, >6 km in diameter) is probably not carcinogenic (IARC, 1988).Wool fiber is classified as a possible human carcinogen (IARCGroup 2B).
Para-aramid fibers
Minimal evidence for skin irritation, none for skin sensitization.Prolonged overexposure to respirable fibrous particles (RFP) has potential for lasting lung damage.
Ceramic fibers
Commercial fibers are too large to inhale, but
RFP can be created when the fiber is abraded or cut. Airborne RFP concentrations from composite machining have been acceptably low. Inhalation studies in rats demonstrated fibrous particle breakdown in the lungs and no carcinogenicity. IARC classifies paraaramid RFP in Group I11 (not classified as a carcinogen).
Skin, eye, or upper respiratory irritation is possible.
be chemically coated or stiff enough to cause irritation by penetrating the skin or tissues of the nose, throat or bronchi. Little has been studied and is known about cured materials being ground, drilled, milled, cut or sanded. Should fragments of fibers be small enough to be respirable, there is concern that a general or fibrous dust hazard to the lungs can occur. Therefore, precautions should be used to minimize exposure. Knowledge of the fiber used should include: length, diameter, aspect ratio and fragmentation propensity. Knowledge of fiber parameters, along with how the composite will be handled, is critical to protect from exposure. Fiber diameter size for respirability is
EPA proposed ceramic fibers, such as refractory aluminum oxide and zirconium oxide, as not classifiable to human carcinogenicity. Also, EPA proposed that aluminum silicate fiber be classified as a human carcinogen. Refractory ceramic fibers are classified as 2B by IARC.
believed to be 3.5 pm or less. Anything larger than that will be removed from the body via nose and throat functions. If the fibers are respirable, the toxic effects may vary significantly. There are low risk fibers (irritants such as fiberglass) and there are significant risk fibers that can result in asbestosis or cancer (such as asbestos fibers). Therefore, each fiber should be assessed on its own toxicological properties. 37.5.3 SOLVENTS
Solvents are used in many aspects of composites manufacturing, from resin formulation to clean-up activities. There are several groups of
836
Safety and health issues
Table 37.10 Health effects of solvents Solvents
CAS nurnber(s)
Ketones
Key notes
Known health effects Mild to moderate skin irritant; moderate to severe eye irritant; if overexposure by inhalation, possible central nervous system (CNS) depression. Irritation of mucous membranes; headache, nausea. Skin contact can cause defatting, dermatitis. Systemic effects only d e r repeated overexposure. If ingested, vomiting may cause acute chemical pneumonitis (lung damage).
PEL-TWA of 1000 pprn and TLVTWA of 500 ppm. ACGIH indicates a TLV-STEL of 750 ppm.
1. Acetone 2-propanone, DMK
67-64-1
2. Methyl ethyl ketone, MEK 2-butanone
78-93-3
3. Methyl isobutyl ketone, MIBK
PEL-TWA of 100 pprn and 108-10-1 Moderate to severe eye irritation; slight to moderate skin irritation; toxic by ingestion and TLV-TWA of 50 pprn and TLV. dermal exposure; may cause CNS depression. STEL of 75 ppm. Overexposure may cause kidney and liver effects.
Few ill effects have been reported. Objectionable odor is reported. MEK has an odor threshold of 0.25-25 ppm. Eye, nose and throat irritation at greater than 200 ppm.
PEL and TLV-TWA of 200 pprn and TLV-STEL of 300 ppm. At the TLV, workers complain of odor, but few ill effects have been reported.
Exposure to high concentrations can cause cardiovascular effects (sensitizationof the cardiac muscle); CNS depression; chronic animal exposures have caused liver and kidney changes.
Chlorinated solvents
1. Methylene chloride Dichloromethane
75-09-2
Vapors below TLV levels; no adverse health responses expected; overexposures may cause possible respiratory irritation due to vapors which could lead to delayed pulmonary edema and CNS depression. Liquid may cause skin and eye irritation.
TLV-TWA of 50 ppm, A2.PELTWA at 25 ppm, PEL-STEL of 125 ppm. Possible human carcinogen. IARC Group 28, NTP Group 2.
2. l,l,l-trichloroethane Methyl chloroform
71-55-6
Drowsiness; overexposures can cause CNS depression which could lead to respiratory arrest; animal studies indicate possible liver and kidney damage.
PEL and TLV-TWA of 350 pprn and TLV-STEL of 450 ppm. MOSH has a recommended exposure limit of 350 pprn as a 15 minute ceiling.
Irritating to the skin, eyes and mucous membranes; stomach pain and cramps; nausea and vomiting; DMF is readily absorbed through the skin and may aid in absorption of other materials; possible link to testicular cancer.
Possible human carcinogen (IARC, Group 2B) PEL and TLV-TWA of 10 pprn (skin).A lethal single oral dose of DMF for humans is estimated to be 10 g.
Other solvents 1. Dimethylformamide, 68-12-2 DMF
2.- N-Methylpymolidone, NMP
872-50-4 Severe dermatitis possible; irritating to skin, severe eye irritant; inhalation of high concentrations can cause headaches, giddiness, mental confusion, nausea, gastric upset, and vomiting; eye irritation.
Maximum airborne levels not established. Manufacturers suggest a TWA limit of 10-100 ppm (vapor), TWA of 5 mg/m3 (mist).
References 837 solvents, ketones and chlorinated, which have dominated the industry. Known health effects of major solvents are outlined in Table 37.10. With increased concern for safety, health and environmental impacts of solvents in the workplace, the use of solvents is decreasing and exposures are being minimized. When choosing a solvent, make certain that regulatory investigation is done to avoid unnecessary rework. It is possible that a solvent of choice may become obsolete due to emission regulations or toxicity concerns.
enforcement’ safety philosophy to control hazards. This encompasses all the workplace tools and techniques discussed. Safety, encompassing health, should be treated as a value. It is not a priority to be raised or lowered with business cycles or management changes. It should be a constant philosophy exhibited each day (Fig. 37.8). If it is a value in the composites industry, then the technology of composite materials can develop naturally, without unnecessary hindrances impeding growth.
37.6 CONCLUSION
REFERENCES
Every industry or business has hazards. The maturity of the industry, its products, and history are all factors in what the hazards or risks are and how they are controlled. There are known potential health and safety hazards in the composites industry, even though it is a young technology compared to mature ones such as steel. As the composites industry develops, processes may change and more specific health and safety issues may be identified. Until then, each composite user has the opportunity to implement the National Safety Council’s ’engineering, education and
American Conference of Governmental Industrial Hygienists. 1997. 1997 TLVs@ and BEIS@, Threshold Limit Valuesfor Chemical Substances and Physical Agents, Biological Exposure Indices. ACGIH. International Agency for Research on Cancer, World Health Organization. 1997. IARC Monographs on the Evaluation of Carcinogenic Risks to Humans. Vol. 68. IARC. Laing, Schmidt. 1992.Accident Prevention Manual for Business G. Industry, Administration G. Programs, 10th Edition. National Safety Council. Laing, Schmidt. 1992.Accident Prevention Manual for Business t3 Indust y, Engineering G. Technology, 10th Edition. National Safety Council. OSHA Instruction CPL 2-2.20B CH-2. 1993. Polymer Matrix Materials: Advanced Composites, Chapter 16. OSHA. Plog, Benjamin, Kerwin, Schonfeld. 1988. Fundamentals of Industrial Hygiene, Third Edition. National Safety Council. Suppliers of Advanced Composite Materials Association. 1996. Safe Handling of Advanced Composite Materials, Third Edition. SACMA. Suppliers of Advanced Composite Materials Association. 1990. Save Your Skin! A Guide to the Prevention of Dermatitis. SACMA.
Although safe and healthful working conditions can be justified on a cold dollars-and-cents basis, I prefer to justify them on the basic principle that it is the right thing to do. In discussing safety in industrial operations, I have often heard it stated that the cost of adequate health and safety measures would be prohibitive and that ‘we can’t afford it.’ My answer to that is quite simple and quite direct. It is this: ‘If we can’t afford safety, we can‘t afford to be in business.’ Admiral Ben Moreell President Jones and Laughlin Steel Corp., 1948
Fig. 37.8 Safety philosophy (Laing and Schmidt, 1992). Used by permission of the National Safety Council, Itasca, Illinois.
ACKNOWLEDGEMENT
Thanks and appreciation to the chemists, industrial hygienists, toxicologists, and safety professionals in the composite industry, particularly those in SACMA’s Environmental, Safety, and Health Committee.
NONDESTRUCTIVE EVALUATION METHODS 38 FOR COMPOSITES Thomas S. Jones
38.1 INTRODUCTION
The development and selection of nondestructive evaluation (NDE) techniques for application to composite materials and structures presents several challenges and considerations that are quite distinct from the considerations given to the similar processes for metallic materials and structures. A principal consideration is the nature of advanced composites as typically layered, anisotropic materials. The materials of interest include fiber-reinforced plastics such as 'fiberglass' and carbon epoxy, as well as some of the more exotic materials such as metal-matrix or ceramic matrix composites. In some cases, naturally occurring composites, such as wood, with its mix of differing density summer and winter growth rings and fibrous structure, may be treated with approaches similar to those used for the man-made composites.The advantages offered by composites are focused on the high strength, low weight properties of typical constituent materials. Yet, if the materials are to exhibit high strength, they must be manufactured as the designer envisioned and they must maintain their integrity in service. N D E presents a technology to help assure the reliability of the materials. Inhomogeneities that may affect the performance of a composite include the concentration
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
of constituents (fiber-resin ratio, resin-starvation, etc.), orientation and distribution of reinforcement, voids, matrix-reinforcement bonding and similar characteristics. The purpose of the NDE is to detect these inhomogeneities and others, including foreign material, fiber breakage, degradation due to moisture, ultraviolet (W)or other reasons, cracks, abrasion, impact damage, fire or excessive heat, etc. This chapter describes many N D E methods that can be used to detect anomalies in composite materials during manufacture and in-service. Although composite materials have been in use on military aircraft for many years, the last few years have seen a rapid escalation in both the number and structural criticality of composite applications entering service. The falling price of composite raw materials has vastly increased the number of applications in non-aerospace industries, including automotive, sports, boating and construction. While the service record for composites has been excellent, they are subject to damage from sources such as overload, hail, lightning, low velocity impact, ballistic rounds and moisture intrusion. Low strain designs, coupled with less critical applications, have made for relatively large acceptable flaw sizes and lax inspection criteria. As composite structures representing a more aggressive design criteria and less tolerant materials (such as some of the ceramic matrix materials) begin to accumulate service hours, we will see a new level of
Visual inspection 839 demands placed on nondestructive inspection techniques to support the in-service maintenance and repair requirements for composites. Several a ~ t h o r s l -have ~ reviewed a variety of nondestructive methods applicable to the evaluation of composites. Several of these methods have, so far, been suitable only for laboratory applications. Some of these inspection methods are adaptations of inspection techniques that have been used for the evaluation of metallic components for years'O. New methods are being developed to address the special needs of composite materials. Many inspection approaches have been automated for production applications; but in-service inspection of composites is frequently performed using conventional manual inspection equipment. Inspection times with manual approaches can be long, but more significantly, interpretation of the inspection results may require special training or familiarization for the inspectors. The long inspection times may restrict inspections to local areas based on criticality, loading, or suspicious areas detected by visual inspections. In recent years, there have been quite a few government sponsored research programs directed at the development of new nondestructive inspection systems providing high throughput for the increasing volume of composite structures in service and/or providing some level of advanced signal analysis which will aid in the recognition and mapping of flaws. As an example, a recent research program resulted in the development of a computer-controlled ultrasonic scanning system for field applicationsll. In this article the capabilities and limitations of some of these typical inspection approaches are considered. 38.2 VISUAL INSPECTION
Some translucent composites can be inspected by transmitted light. Inhomogeneities such as voids, delaminations or inclusions can be detected by this straightforward approach in materials such
as nonpigmented glass reinforced plastic (GIG'). Stress whitening in GRP, which results in a loss in optical transparency, may result from fiber-resin debonding or resin cracking. Another visual inspection approach, which has been considered for the detection of impact damage, is to paint the composite surfaces with a paint containing micro-encapsulated dye. When crushed by the impact, the dye is released and reveals the location of the impact. Dyes can be formulated either for visual observation of a color change or using UV excited fluorescence in the dye. This method provides good indications to locate suspicious areas for further inspection so that the extent of the impact damage can be determined. Visual aids, such as magnifiers, borescopes, television cameras, etc., can be used to enhance the detectability of surface-related damage. Small diameter, flexible borescopes permit the inspection of inaccessible areas without teardown of components. The most common visual inspection is exemplified by the typical walk-around visual inspection of aircraft. These inspections certainly apply to much of the composite structures currently in use. Many of the more severe conditions associated with composites are visually detectable. Punctures, surface ply delaminations, scratches, gouges and heat damage can frequently be detected by visual inspection. Disbonds between a composite skin and some substructure may also be detectable in some cases as a blister in the skin, an edge separation, or a distortion in the geometry of the component. This inspection is clearly valuable, but one cannot expect to detect all of the forms of damage which may be present in a composite structure. Further, if damage is detected, it is important that other tools are used to assess the extent of the damage, since the subsurface damage may drastically exceed the surface detectable damage. One common example is low velocity impact damage. S t u d i e ~ l ~on , ' ~this form of damage, which may result from hail, runway debris, or ground handling abuse, have shown
840 Nondestructive evaluation methods for composites that substantial subsurface damage can result with little or no surface detectable condition.
disadvantages as manual tapping with the additional disadvantage of increased cost.
38.3 TAPTEST
38.4 ULTRASONIC METHODS
The tap test method, using either a coin or a special tap hammer continues to be a common in-service inspection tool despite the availability of less subjective inspection tools. The use of this method has persisted for several reasons. The most obvious is that it does not require sophisticated or expensive equipment. Another is that many of the composites in use consist of thin laminates in low strain designs. This combination yields critical flaw geometries which are fairly large and close to the surface, two conditions necessary to the successful use of the tap method. Under these conditions, the tap method can be a useful tool for the detection of problems in relatively large areas of laminate, particularly where the substructure of the tested skin is relatively consistent. The tap test method is sensitive only to laminar type flaws, such as delaminations or unbonds, and relies on the different acoustic resonance of the loose upper layer compared to the surrounding material. The tap method suffers from subjective interpretation (particularly with complex geometries), variable application, declining sensitivity with flaw depth and an inability to calibrate effectively for either flaw size or depth. The more recent applications of composites in thicker laminates and more highly loaded designs make this approach inadequate in many cases. Several attempts have been made to instrument the tap test by providing a machine-type tapper and instrumentation to interpret the signals. The mechanical tappers offer the advantage of improved repeatability in terms of tap impact and location. The instrumentation developed centered on spectral analysis of the detected audio signals. These instrumented tappers have not made a significant impact on the composite NDE instrument market. They suffer from many of the same
38.4.1 ULTRASONIC THROUGHTRANSMISSION TESTING
Ultrasonic inspection makes use of €ugh frequency (above 20 kHz) mechanical vibrations. Typical ultrasonic frequencies for inspection of composite materials are in the megahertz range. Ultrasonic signal strength in a material is reduced both by attenuation and by reflections from interfaces. Large reflection signals are obtained when ultrasound is directed from one medium to another with the two media having very different acoustic impedances. For example, more that 99.9% of an ultrasonic wave is reflected from a metal-air interface as might be encountered with a crack in a metal. The ultrasonic signal is typically introduced in a pulsed mode. The inspection geometry can be through-transmission, with the receiver ’listening’ for the ultrasonic pulse on the opposite side of the component from the transmitter, or a one-sided pulse-echo test, in which a single transducer first transmits a pulse and then ‘listens’ for the reflections of that pulse from the interfaces of the inspection object. Ultrasonic through-transmission test systems measure the signal strength of a pulse of ultrasonic energy transmitted through the structure or material under test. Locations where there may be a delamination or a foreign material, for example, will show a reduced exit ultrasonic intensity. This test approach is probably the most commonly used production inspection method for composite structures. It is relatively easily automated, provides approximately constant sensitivity to flaws, regardless of their depth within the structure and is fairly easy to interpret. Sophisticated computer controlled inspection systems, such as the McDonnell Douglas Automated Ultrasonic Scanning System (AUSS)14, shown in Fig. 38.1, are
Ultrasonic methods 841
Fig. 38.1 McDonnell Douglas AUSS ultrasonic system.
common in aircraft production facilities. These systems scan the composite component while collecting and recording the ultrasonic pulse transmission amplitude. The method is sensitive to most flaws which are planar in nature and lie roughly parallel to the surface. Since this describes most of the common flaw types for layered composite structures, the method is very appropriate. Where automated testing systems are used, the test can be performed very efficiently and can yield a recording, or map of the inspection results, called a C-scan. An example of a through-transmission C-scan showing a delamination caused by a foreign material inclusion is shown in Fig. 38.2. Generally, the method requires access to both sides of the part and alignment of ultrasonic search units on opposite sides of the part. These factors drastically restrict the usefulness of this approach for in-service inspection, since in many cases, access to both sides is not available or not practical. Further, the speed and recording advantages of auto-
mated systems is usually not available for field inspection. This is, however, changing. Many of the military depot inspection facilities now I
.
, +
,
, 1
..
,
, .- -, . .
,
.
, -,
I
I
I 8
4
-
Fig. 38.2 Computerized through-transmission C-scan showing foreign material.
842 Nondestructive evaluation methodsfor composites
have automated ultrasonic inspection systems for composite inspections. Further, inspection systems such as the Automated Real-time Inspection System (ARIS)15provide a semiautomated through-transmission inspection and data recording capability for some on-aircraft inspections. A 9 kg (20 lb) yoke provides the mechanism to place an ultrasonic transducer on either side of the field inspection object. One-sided pulse-echo testing can also be accomplished. 38.4.2 ULTRASONIC PULSE-ECHO TESTING
The ultrasonic pulseecho inspection approach typically uses a single search unit as both the transmitter and receiver. The approach requires access to only one side of a material or structure to be tested. Flaws are detected by monitoring the time of arrival and/or the signal strength of returning echoes. Delaminations cause the returning echo to arrive prior to its expected travel time from the round trip from the front surface to far surface and back to front surface. This method is commonly used for in-situ inspection since the single transducer approach usually makes it simpler to apply in a manual test. On the other hand, the more complex signal patterns and more critical sound entry angle, make this test more difficult to automate than the throughtransmission test. Even so, computer automated pulse-echo inspection systems have gained popularity in recent years. An example is the Large Area Composite Inspection System (LACIS). This semiautomated pulse-echo inspection system is useful for in-service inspection of composites. The LACISll (Fig. 38.3) uses a reciprocating arm holding two to four transducers to perform pulse-echo inspection of large composite components quickly. Inspection speeds in excess of 9.3 m2/h (100 fP/h) have been reported with this hand-held scanning device. An advantage to the pulse-echo systems is that flaws at multiple depths can be distinguished from one another. The pulse-echo
Fig. 38.3 McDonnell Douglas LACIS portable ultrasonic scanner.
C-scan shown in Fig. 38.4 shows the multiple delaminations associated with impact damage. In this scan, the delaminations show as gray to black areas in which the darkness of the indication reveals its depth. The darker indications are closer to the surface. The football shaped delamination of the far surface ply is seen as a gray shape only slightly darker than the normal background thickness. The pulse-echo method is also advantageous in that it offers increased sensitivity to foreign material inclusions associated with the manufacturing process for laminated composites. Many expendable paper and plastic materials are used in handling and transporting
Fig. 38.4 pulse-echo time-of-flight C-scan showing
impact damage.
Ultrasonic methods 843 uncured composite materials. These materials have at times found their way into a composite laminate and are cured in place. Some of the plastic materials are particularly problematic in that they may bond into the laminate and go undetected by the through-transmission inspection. Fortunately, however, most of these materials offer a reflected signal strength sufficient to be detected by pulse-echo ultrasonics.
surface loading on the ultrasonic probeI8.This can be seen as a phase, amplitude, or resonant frequency shift in the ultrasonic element. This method is particularly useful in complex bonded structures where access limitations restrict the use of ultrasonic through-transmission testing and the complex internal reflections make pulse-echo signals difficult or impossible to interpret. 38.4.5 ULTRASONIC CORRELATION
38.4.3 ULTRASONIC POLAR BACKSCATTER
This inspection approach makes use of slightly angled ultrasonic beams to detect flaws in composite materials. The angle orientation frequently allows one to associate flaws with a particular ply orientation. The method is particularly useful for characterizing matrix cracking within the composite plies. It is also sensitive to linear voids or porosity. It has been successfully used"j to characterize the various levels of damage in impacted laminates. The approach has also been used to measure directional velocity variations and elastic properties in composite^'^. From these measurements, the anisotropic elastic properties can be calculated. At the current time, this inspection method remains primarily a laboratory tool. The angulating mechanism would be somewhat difficult to manipulate and control in the field and the inspection times can get long, since multiple scans at a variety of orientations are required. Further, the composite designs currently in use have not required the detailed level of inspection provided by this method. 38.4.4 ULTRASONIC RESONANCE
This one-sided ultrasonic inspection method detects laminar discontinuities within composites or bonded structures by setting up a continuous ultrasonic wave within the material and sensing the mechanical stiffness, or impedance of the material. A delamination or disbond reduces the surface normal stiffness of the material and this, in turn, reduces the
Ultrasonic correlation offers a novel approach to the ultrasonic inspection of highly attenuative materials19. It has proven effective in the evaluation of some of these materials where conventional pulsed ultrasonic systems have experienced difficulties in penetrating the material. The correlation approach achieves this increased sensitivity using a continuous wave, cross-correlation technique which enhances the sensitivity of the test but sacrifices inspection speed. A block diagram of a continuous wave correlator is shown in Fig. 38.5. The principal advantage of the correlator is that it provides substantial enhancements in the signal-to-noise (S/N) ratio of the ultrasonic signals. This S/N enhancement is achieved through a modification of the time scale required to produce an A-scan trace. Since the correlator uses continuous generation and accumulation of the ultrasonic signals, the maximum possible efficiency of data accumulation can occur. The S / N enhancement allows the correlator to produce usable A-scan traces where conventional pulsed ultrasonic systems are unable to produce a recognizable pattern. For example, the correlator has been used to characterize polytetrafluoroethylene (PTFE, Teflon@),a material very difficult to penetrate with conventional ultrasonic techniques. The results have been obtained by measuring ultrasonic velocity as the temperature of the PTFE sample changed. Figure 38.6 shows a velocity versus. temperature profile with an excellent indication of a change in the slope of
844 Nondestructive evaluation methods for composites 7
Transmit
Test Specimen
Noise
I
I
Fig. 38.5 Block diagram of ultrasonic correlator.
the curve at the phase transition temperature of 30°C. The correlator has been shown effective for the inspection of thick and highly attenuative composites. Effective signals have also been obtained through over 30 cm (11.8in) of cured wood and through wood/plastic laminates, such as countertops.
1340
(dS)
i
1240
1
1220
I
1320 1300 Velocity i280
1260
20
25
30
35
40
Temperature ('C)
Fig. 38.6 Ultrasonic velocity transition in PTFE measured with correlator.
38.5 X-RADIOLOGY
X-ray imaging relies on the differential absorption or scattering of the X-ray photons as they pass through a material. Flaws which either allow more X-ray photons to pass or which absorb or scatter the photons can be imaged if the effect is significant. X-ray inspection systems are usually sensitive to changes which result in an apparent change of at least 1-2% of the material thickness or density. Composite flaws tend to lie between plies and present a very small apparent thickness change, particularly for thick laminates. However, the low density of most composite materials permits the use of low energy X-rays which help to enhance the sensitivity*O.X-rays can be used to detect porosity and matrix cracks as well as some foreign materials. More highly loaded applications and/or the use of more brittle matrix materials, both of which seem to be on the horizon, will make the detection of some of these conditions much more critical than they
X-radiology 845 are today. While carbon fibers are not generally imaged by X-rays, boron and silicon carbide fibers are typically deposited on a tungsten filament and this filament can be imaged in X-radiographs. This allows one to detect fiber fractures and determine fiber orientations and placements, so long as the number of plies does not get too large. X-radiography is particularly useful for the detection of honeycomb core defects in bonded sandwich assemblies. The low density and thin composite skins usually provide minimal interference for the X-rays to image the honeycomb core materials. Core defects such as blown core, crushed core, condensed core, fatigued, corroded or cut core and foaming adhesive voids can be detected by radiographic methods. It is also frequently possible to detect water intrusion into the honeycomb core using X-ray methods. Significant progress has been made in recent years in the use of non-film electronic imaging systems for X-rays. These systems allow the real time viewing of X-ray images and can drastically reduce the demand for Xray film. While video imaging systems do not
typically possess the resolution of X-ray films, the use of very small X-ray sources and geometric magnification techniques have permitted the recovery of much of this sensitivity. Figure 38.7 shows a magnified image of a honeycomb core area. This 'bee's eye view' of the honeycomb core reveals the onset of damage in the cell wall. In practice, one must find a suitable compromise between spatial resolution and area of coverage. The cost advantages of this approach have made it very attractive. Many military aircraft maintenance facilities either have or are acquiring real-time X-ray inspection equipment. 38.5.1 X-RAY BACKSCATTER IMAGING
Conventional radiographic inspection techniques rely on the attenuation of a beam of penetrating radiation to form an image of a part. At low X-ray photon energies typically used with composites, a large portion of that attenuation is due to Compton scattering. This suggests the use of scattered radiation to analyze the inspection object. One approach utilizing a novel, X-ray backscatter camera
Fig. 38.7 Microfocus X-radiograph of honeycomb cell showing onset of damage.
846 Nondestructive evaluation methods for composites Lamina
Fig. 38.8 X-ray backscatter system geometry.
insulator and a rubber liner, as shown in Fig. 38.9(a).Figure 38.9(b) is the X-ray backscatter inspection result for a single location. The illustration shows the relative X-ray backscatter intensity versus thickness of the inspection sample; depth through the sample can be correlated with detector number. Region A represents the near-surface graphite-epoxy region. The sharp dip in response at B represents the decreased X-ray backscatter intensity typical of an air gap or delamination. The insulator response is given in region C; note the differing response slope for this lower density material, as compared to the response for the second graphite-epoxy layer shown in region D. The response from the rubber deeper in the assembly is shown in region E. The X-ray backscatter signal contains quantitative information about variations in density as caused by changes in material or delaminations and the location of such variations within the depth of the material. These characteristics, coupled with the one-sided inspection feature, represent advantages for the inspection of composite assemblies.
employing a slot instead of the more common pin-hole approach has been developedz1.This technique is illustrated in Fig. 38.8. The backscattered X-rays are detected by an array of scintillation detectors. The backscatter intensity information is obtained as a function of detector position; this in turn, relates to different depths in the inspection object. One interrogates the entire object thickness at one 38.5.2 COMPUTED TOMOGRAPHY time (composite structures as thick as 7cm (2.7 in) have been inspected). The inspection Tomographic inspection systems, developed results provide indications of material density originally for medical diagnostics, have been changes by a change in the slope of the inten- further developed for applications in the A tomographic sity/position graph and of voids or foreign industrial inspection material by an abrupt change in backscatter image looks like a slice taken across the intensity. The inspection method is particu- inspection object. A collimated single pencil larly useful for the inspection of laminated beam or a flat fan beam of X-rays is directed at structures such as pressure vessels and rocket the sample. The sample is shifted and rotated motor cases. In some of these designs, ultra- relative to the X-ray beam while the X-ray sonic inspection approaches are ineffective or beam intensity is measured by individual impractical and the backscatter X-ray imaging detectors (typically arrays of detectors) at each approach offers a potential solution. Tight position and rotation. The measurements perdelaminations, with gaps less than 50 p m in mit a computer reconstruction of a density map of the inspection object. One advantage of width, can be detected. An example of an X-ray backscatter inspec- tomographic inspection is the extremely good tion is shown in Fig. 38.9. The sample was a contrast sensitivity; variations of less than graphite-epoxy composite structure as used in 0.1% are detectable. The major advantage of a rocket motor case. The assembly included computed tomography is that the resultant several layers of graphite epoxy composite, an image shows all the variations across the
X-radiology 847 BACKSCATTER X-RAYS
INCIDENT X-RAYS
GRAPHITE EPOXY
(4
r
98.
54.
43.
32. 21.
IO.
(b)
1 1 1 1 1
2
3
4
5
8
7
8
9
10 11
12 13 14 15 18 17 18 I9 20
D€ECTORNWBW
Fig. 38.9 X-ray backscatter examination results (a) composite sample configuration and (b) system
response.
image slice in the true three-dimensional perspective. A typical radiograph is a two-dimensional projection of a three-dimensional object. The tomographic result retains the three-dimensional nature of the image. A computed tomography image of a tree, for example, would appear similar to the visual image if one sawed through the tree at one location. The tomographic image would show the differing density of summer and winter growth and the presence of knots, voids, etc.
For composites, tomography offers similar advantages. The tomographic image shown in Fig. 38.10 shows variations in density across the width of the composite material. 38.5.3 NEUTRON RADIOGRAPHY
X-rays are attenuated as a function of the density of the material through which they pass. If X-ray attenuation is plotted as a function of increasing atomic weight of the attenuating
848 Nondestructive evaluation methods for composites
I
. .' ,
1
,
radiography. Recent advances in moderating and collimating accelerator fast neutron beams have led to significant improvements in portable system thermal neutron beamsz6.A typical thermal neutron image of a composite structure, as obtained with the transportable inspection system shown earlier, is shown in
I
Fig. 38.10 Computed tomogram of graphite silica composite tube.
ap
material, one finds a monotonically increasing function. The attenuation of thermal neutrons", plotted similarly, shows a generally random pattern in which there is high attenuation for several light materials, hydrogen, boron and lithium in particular and relatively For bonded Fig. 38.11 Photograph of transportable thermal low attenuation for most structures, the high sensitivity of neutrons to neutron radiographic system. hydrogen means that thermal neutron radiography can display images of adhesives, water, or corrosion in a metal bonded assembly more readily than X-ray techniques. For composites, neutron techniques offer promise for the detection of variations in the organic matrix materials and moisture take-up. Advances in transportable neutron sources have been made including source systems for depot and field inspections. A photograph of one system= is shown in Fig. 38.11. The neutron source is a Cockroft-Walton accelerator that produces neutrons by accelerating deuterons on a tritium target, a D-T reaction that produces neutrons of about 14Mev in energy. These neutrons are slowed, or moderated, in a low atomic number assembly and Fig. 38.12 Thermal neutron radiograph of a bonded collimated into a beam for thermal neutron fiberglass-epoxy structure.
'
I
Eddy current testing 849 A major neutron radiographic facility for the inspection of military aircraft has been installed at McClellan Air Force Base. A maneuverable, crane-based system permits near real-time neutron inspection of wing and empennage assemblies on the aircraft. Film techniques are used to inspect lower doors. These on-aircraft inspection systems employ the radioisotope 252Cfas the neutron source. The system has emphasized the inspection of metal bonded assemblies to detect water intrusion and corrosion damage. A nuclear reactor facility is available at this facility to inspect components removed from the aircraft. Consideration is being given to replacement of the californium source with an accelerator source of neutrons.
One area where acoustic emission testing has found acceptance is in the detection of moisture and corrosion in honeycomb assemblies. If an aluminum honeycomb assembly containing water is locally heated in the vicinity of the water, the increased vapor pressure will force the water to move through the leakage path along which it entered the honeycomb cell. The motion of the water through the leakage path can be detected using acoustic emission equipment. 38.7 ACOUSTO-ULTRASONICS
The acousto-ultrasonic, or stress wave factor, test method uses an ultrasonic transducer to inject a simulated acoustic emission pulse into the material under test. The wave form of the pulse is then monitored a fixed distance away 38.6 ACOUSTIC EMISSION with an acoustic emission transducer. Damage Acoustic emission testing involves the detec- to the material will affect the manner in which tion of elastic energy that is spontaneously the wave is transmitted through the materialz8. released by materials when they undergo The stress wave factor will be affected most by deformationz7.For example, when a piece of discontinuities which impede the motion of wood is stressed, audible cracking noises can the stress wave through the material. The be detected just before the wood breaks. The method can be applied in a through transmissame phenomena occurs in other materials, sion mode, but is generally used with both including composites. The detected signals are transducers on the same side of the part; one frequently in the ultrasonic rather than audi- monitors the transmission of signals parallel to ble region. the surface. In this mode the method is most Acoustic emission testing monitors the sensitive to matrix cracking and laminate material under test for spontaneous noise gen- porosity. It is also sensitive to fiber breakage erated under load. It has been used extensively and, particularly in the through transmission in the proof testing of fiberglass pressure ves- mode, delaminations. The method has sufsels and beams. It has also been used to fered from problems in obtaining consistent monitor and characterize damage growth results. Some of these problems have been mechanisms in composites under cyclic load- solved or reduced and the method shows ing. Acoustic emission testing is capable of potential for evaluating the severity of matrix detecting and characterizing matrix cracking, degradation in composites. One approach has delamination and fiber breakagez7.It is used been to use completely non-contact generation extensively in the testing of composite pressure and detection of the signalsz9. vessels and in the evaluation of some structures, such as 'cherry picker' booms; however, 38.8 EDDY CURRENT TESTING the method has not found particular application in the aircraft inspection arena for these Eddy current methods rely on the principles of flaws. This may be partly due to the need to magnetic induction to interrogate a material have the component under load during testing. under test?. Current loops, or eddy currents,
850 Nondestructive evaluation methods for composites are induced in a conducting material by a varying magnetic field. While most resin matrix materials are very poor electrical conductors, some of the carbon fibers are relatively good conductors. In order to form current loops, multiple fibers must make electrical contact with one another at various places along the length. Fortunately, the use of carbon fiber tows of 3000 to 10000 fibers or more clearly promotes this. Eddy currents can then be used to monitor fiber orientation since the current loops are not circular, but greatly elongated along the fiber direction. Eddy currents can also be used to detect fiber breakage. Eddy current techniques have been demonstrated to provide effective results for many of the damage mechanisms, including impact damage and fatigue damage, in carbon fiber composite^^^. Even so, eddy current techniques have not received a great deal of acceptance to date for production or in service inspection of composites.
(T
= Stefan-Boltzmann constant = 5.7 x lo4
T
= absolute temperature, degrees Kelvin
(W/m2K4)
The emittance in a particular differential wavelength band is governed by the Planck distribution criterion. Important criteria for practical infrared imaging are the spectrum of the emitted energy and the wavelength of maximum emittance. These are given by the single temperature evaluation of Planck's Law and Wien's Displacement Law, respectively. Figure 38.13 shows a family of such energy distribution curves. For the practical case of a body at a temperature of 25°C (77"F, approximately room temperature), the wavelength range of peak emission is in the range 9 - l o p . This turns out to be a very useful wavelength range. Various components of our atmosphere, notably water, absorb a great deal of the emitted infrared energy. However, there are two fairly transparent windows through the atmosphere: one between 3-5 p and the 38.9 INFRARED THERMAL TESTING other at about 8-14 p.Specific infrared Infrared thermography has been gaining detector elements have been developed which rapidly increasing acceptance as a nondestruc- respond well to each of these wavelength tive evaluation tool for composite (as well as bands. As seen in Fig. 38.13, the intensity of many other) structures. All bodies above the infrared emittance within such a band can be temperature of absolute zero emit electromag- used to indicate the temperature of the object netic radiation by virtue of the motion of the surf ace. constituent atoms. The spectrum and intensity 0.0005 of the radiation depend on the temperature and nature of the surface. When a surface is -T=50 'C heated, there is an increase in energy of the 0 . m. atomic particles leading to a corresponding increase in temperature and emitted f 0.0003en erg^^^,^^. The wavelength-independent rate -e of emission of radiant energy per unit area is m I o.ooo2 governed by the Stefan-Boltzmann law:
-
w = &UT4 where: W = rate of emission, radiant energy per unit area E = emissivity (ratio of emittance of the surface relative to a black body)
o.ooo1-
07 0
2
4
0
8
10
12
14
I0
18
Wavelength (pm)
Fig. 38.13 Spectral radiant emittance distribution at three surface temperatures.
Laser shearography/holography Typical thermographic applications involve the introduction of a controlled thermal load on the object of interest. Variations in the thermodynamic properties of the object then produce surface temperature patterns which can be detected with an infrared imaging system. Composite materials typically offer an excellent combination of thermal properties for the useful application of infrared thermography. Thermographic methods have been used to display object temperature differences for a variety of inspection applications. Many of the inspection systems provide a television-type image display. Typical sensitivityof the instruments under ideal conditions is 0.1"C (0.2"F). Thermal patterns can be produced by heating the inspection surface and monitoring the surface of the part for relatively hot spots caused by the blockage of the thermal flow away from the surface by a laminar flaw. This method is particularly sensitive to flaws near the inspection surface; the sensitivity reduces rapidly with flaw depth3. The thermal images can also be produced by heating the far side of the object and monitoring the surface for cold spots where the heat flow toward the inspection surface in impeded by the flaw. This 'through-transmission' approach somewhat improves the depth sensitivity, but the method is still more sensitive to near side flaws. Impact damage often results in matrix damage near the surface of a composite material. Figure 38.14 shows an infrared image of an impact damaged graphite epoxy composite sample. The object was heated and imaged on the same side. The diamond shaped indication is typical of the delamination pattern produced in a cross-plied laminate. Infrared thermography has been demonstrated to be an effective tool for evaluating many glass fiber composite structures used in marine applicati0ns353. Infrared thermography has been gaining increased acceptance in recent years as a viable nondestructive testing tool. A significant number of effective applications has been
851
Fig. 38.14 Thermogram of impact damage in graphite epoxy composite.
developed for composite materials and structures. This increased acceptance has been assisted by improvements in the infrared imaging technology, development of improved heat applications techniques and significant efforts to model and understand the heat flow characteristics in anisotropic media. Effective applications have been identified for a wide range of materials and applications. Thermography has been used for aerospace laminates of graphite epoxy as well as for marine applications of fiberglass up to around 8 cm (3 in) thick. In many instances for materials which are most difficult for more traditional nondestructive tests, such as ultrasonics, infrared thermography has produced rapid, effective results. An example is foam core fiberglass boat hull structures. While this structure poses substantial difficulties for ultrasonic techniques, infrared thermography has produced rapid, effective results5. 38.10 LASER SHEAROGRAPHY/ HOLOGRAPHY
A hologram is an interference pattern that can be used to reconstruct the optical wavefront that originally emanated (due to reflection or transmission) from an object. The hologram is
852 Nondestructive evaluation methodsfor composites
formed by the superposition of two wavefronts, an object beam and a reference beam, on a suitable recording material, such as photographic film.When properly illuminated by the reference beam, an observer looking through the developed hologram sees a virtual image of the original object. When the real object undergoes a small displacement over part of the surface, due to stressing by thermal or mechanical means, a variation in the relative phase of the wavefronts will be produced and a fringe pattern can be observed. Laser illuminationprovides the spatial and temporal coherence in the light beam to permit the observation of the fringe pattern, a technique called holographic interferomeW7.An example of a holographic interferometry inspection of a composite tube to reveal impact damage is shown in Fig. 38.15. Laser interferometric holography techniques were evaluated for composites testing in the 1960s. In fact, several composites and bonded assembly testing systems were constructed. In spite of the use of massive optical
benches and elaborate vibration isolation systems, however, the vibration problems of working in a production environment (much less an in-service environment) were not satisfactorily solved. The holographic inspection approach is sensitive to very small amounts of relative motion in the part surface. Typically a part is tested by taking an image of the part at rest, then stressing the part surface, either with a small amount of heat or a mild vacuum. This light stressing causes the area of the surface over the flaw to deform slightly more than the surrounding material. Displacements as small as one-quarter wavelength of the laser light being used cause a fringe to appear in the interference pattern. Unfortunately, unwanted test object motion of an equally small amplitude will also cause fringes. The development of phase locked loop holographic systems has gone a long way to eliminate the problems associated with low frequency vibration. Holographic inspection systems which use video imaging systems and 'develop-in-place' reference holograms are currently being used.
Fig. 38.15 Holographic interferogram of impact damage in a composite tube.
Conclusions 853 Another development in the area of video holographic systems, electronic shearography, has recently provided even greater immunity to test object vibration and motion problems, making possible the rapid scanning of large areas of composites on aircraft. Electronic shearography uses no film and compares a live video image with a stored video image to produce interference fringes. Conventional interferometric holography interferes two holograms of a component to produce the interference pattern. Each hologram is produced by recording the speckle interference between a coherent reference beam and the coherent object beam which illuminates the surface of the test part. Deformations between the two exposures produce the interference fringes. Phase-locked holography uses the same basic technique, but uses the diffuse reflection of an unexpanded coherent beam shone on a small portion of the test object as the reference beam. This approach minimizes the influence of full surface motions, typical of environmental vibration and is sensitive to differential surface displacements typical of defect indications. Shearographf8, on the other hand, uses no separate reference beam. Rather, the returning object beam is doubly imaged, with one of the images slightly shifted or 'sheared' relative to the unshifted image. Thus, the interference pattern does not reveal bulk surface motion, but only the degree of differential motion of the surface along the direction of the shearing. This makes shearography particularly well suited to many production and depot environments because of its relative immunity to vibration problems39.An example of a shearographic inspection is shown in Fig. 38.16. In this case the inspection object is a sandwich panel consisting of graphite-epoxy laminate skins and a foam core. The indication arises from a delamination created by a pull-out in the skin-to-core bondline. The side-by-side pattern of concentric rings is characteristic of the shearography technique.
Fig. 38.16 Electronic shearography image of graphite*poxy/foam core panel with edge delamination. 38.11 MICROWAVE TESTING
The use of microwave energy to interrogate and characterize composite materials has been investigated for some years. Microwaves are very sensitive to small changes in the dielectric properties of low conductivity composites, such as glass and aramid fiber composites. Microwave techniques have also been investigated for applications to composites with higher conductivity fibers such as graphite fibers. Successful results have been reported for the measurement of fiber content and orientationgo, material thickness41and porosity content42. To date microwave techniques remain largely laboratory tools and have not gained great acceptance as production evaluation techniques. However, these techniques offer excellent sensitivity to conditions, such as matrix porosity and cure state of the matrix, that are difficult to establish by more conventional nondestructive testing techniques. Increased applications for these techniques can be anticipated in the future. 38.12 CONCLUSIONS
A wide variety of nondestructive testing techniques are applicable to the evaluation of
854 Nondestructive evaluation methodsfor composites various composite structures. The selection of the most appropriate test technique for a particular application can be a formidable task. The nondestructive inspection methods most widely used in industry for composite inspection, ultrasonics and radiography, have much to recommend them in that many different types of discontinuities can be detected and characterized. Shearographic and infrared inspection methods present the attractions of noncontact, large area coverage and good sensitivity to discontinuities close to the inspected surface. These techniques are finding a rapidly growing list of applications where the more traditional tools encounter difficulties or are
too cumbersome to apply. Delaminations in the order of 1 cm (0.4 in) in size can usually be detected by shearography and even smaller, near-surface discontinuities can be detected by infrared imaging. Both methods offer a television image display which offers straight forward interpretation. Both methods also offer good sensitivity in low density material systems which are frequently difficult to evaluate using the more traditional techniques. The overall summary of inspection methods as related to composite inhomogeneities includes many inspection methods. Table 38.1 provides a sorting approach that can be used for the selection of inspection methods for par-
Table 38.1 Summary of applicability of NDE methods
Flaw Type Porosity Foreign material Shallow delamination Deep delamination Matrix cracks Fiber breaks Impact damage Skin/skin disbond Skin/core disbond Crushed core Condensed core Blown core Core node disbonds Water intrusion Corroded core Fatigued core Foam adhesive voids Bondline adhesive voids
3 2
1 2 1 1
1 1 1 1
3 3 3
2 2 3
1 1 1 1
1 2 2 3
2 2 2 3 1 2
2 1 3 2 1 2
1
2 2 1 1 1 1 1 1 1
3 2 2
2
1
2
1
3 2 2 3 1
1 2 1 2 2 2 3 1 1 1 1 2 2 1 2 2
2 1 1 2 1
2
1 2 1 1 2 2 1 2 3 2 1 1 1 2 3 2 1 2
Key: 1. Good sensitivity and reliability. Good candidate for primary method. 2. Less reliability or limited applicability.May be good supplementarymethod. 3. Limited applicability. May provide some useful information.
1 2
2 2
3 3 2 2 3 2 2
1 1
2 3
2 1
2
2 2 1 2 2 3
1 1
2 1 2 3
3 1 2
1 1 1 2
1 1 1 2
2
2
1
2 1 1
References ticular comDosite discontinuities. While this table cannot be considered definitive, it does give general guidance to the applicability of the various technologies to various material conditions. In any specific application, a large number of considerations will influence the selection of nondestructive testing technologies. It is hoped that the discussions provided in this chapter will offer some assistance in the selection and application of nondestructive testing methods. However, final selection should be based on the evaluation of a greater number of parameters than can be considered here. Experience, training, experiment and clear understanding of the inspection objectives are required to develop effective nondestructive testing applications.
855
1985. 11. Large Area Composites Inspection System-I,
USAF Contract F33615-91-C-5664, Wright Research and Development Center, WrightPatterson Air Force Base, Contract with McDonnell Douglas Corporation, 1991. 12. Smith, B.T., Heyman, J.S., Moore, J.G., Cucura, S.J. and Freeman, S.M., Correlation of the deply technique with the ultrasonic imaging of impact damage in graphite/epoxy composites, Review of Progress in Quantitative Nondestructive Evaluation, Vol. 5B, New York Plenum Press, 1986,1238-1244. 13. Gros, X.E., Low energy impact detection on carbon fiber reinforced materials, Materials Evaluation, 1995,53,373-381. 14. Jones, T.S., Inspection of composites using the automated ultrasonic scanning system (AUSS), Materials Evaluation, 1985,43, 746-753. 15. Jacobs, B., Hamlin, D., Peterson, R. and Spinks, R., Composites in-service inspection system producibility, Air Force Report AFWAL-TR 88 REFERENCES 4218, Wright Research and Development 1. Prakash, R., Non-destructivetesting of composCenter, Wright-Patterson Air Force Base, OH, ites, Composites, 1980,11(4),217-224. 1988. 2. Scott, LG. and Scala, C.M., A review of non- 16. Blodgett, E.D., Miller, J.G. and Freeman, S.M., destructive testing of composite materials, NDT Correlation of ultrasonic polar backscatter with International, 1982,15(4), 75-86. the deply technique for assessment of impact 3. Reynolds, W.N., Nondestructive testing (NDT) damage in composite laminates, Review of of fiber-reinforced composite materials, SAMPE Progress in Quantitative Nondestructive Quarterly, 1985, 16, 1-16. Evaluation, Vol. 5B, New York: Plenum Press, 4. Shuford, R.J., Hinton, Y.L., Murray, T.J. and 1986,1227-1238. Brockelman, R.H., Advanced N D E techniques 17. Bar-Cohen, Y., Mal, A.K. and Lih, S.S., NDE of for composites, Technical Paper EM85-113, composite materials using ultrasonic oblique Society for Manufacturing Engineers, 1985. insonification, Materials Evaluation, 1993, 51, 5. Bar-Cohen, Y., NDE of fiber-reinforced compos1285-1296. ite materials - A review, Materials Evaluation, 18. Botsco, R.J., New methods for nondestructively 1986,44,446-454. evaluating airframes and jet engines, Air 6. Brahey, J.H., Inspection and repair of composTransport Association Forum, Long Beach, CA, ite aerospace structures, Materials Evaluation, September 8-11,1980. 1986,44,1513,1531,1539. 19. Kishoni, D., Rosen, M., Berger, H. and Cheng, 7. Jones, T.S. and Berger, H., Nondestructive evalY.T., Signal to noise enhancement by an ultrauation methods for composites, International sonic cross-correlation system, IEEE Ultrasonics Encyclopedia of Composites, (S.M. Lee, ed.), 1991, Syrnp., Chicago, Oct., 1988. VOl. 4, pp. 3749. 20. Jones, T.S., Polansky, D. and Berger, H., 8. Seidl, A.L., Inspection of composite structures Radiation inspection methods for composites, Part I, SAMPE Journal, 1994, 30, July/Aug NDT International, 1988,21,277-282. 38-44,1994. 21. Berger, H., Cheng, Y.T., Jones, T.S. and Polansky 9. Seidl, A.L., Inspection of composite structures D., An electronic imaging technique for onePart 11, SAh4PE Journal, 1995,Jan/Feb, 31,4248. sided X-ray inspection, Proc. 1989 ASME 10. Boyer, H.E. and Gall, T.L., (eds), Metals Pressure Vessel and Piping Conference, NDE, Vol. Handbook, Part IV,Section 33, Nondestructive 5, (R.D. Streit, ed), pp. 89-93, Amer. SOC.Mech. Testing, Amer. Soc. Metals, Metals Park, OH, Eng., New York, July 1989.
856 Nondestructive evaluation methods for composites 22. Bryant, L.E. Radiography
and McIntire, P. (eds.), and Radiation Testing, Nondestructive Testing Handbook, Vol. 3, Radiography and Radiation Testing, Amer. SOC. Nondestr. Testing, Columbus, OH, 1985. 23. ASTM E-1441, Standard Guide for Computed Tomography (CT) Imaging, Amer. SOC.Testing Materials, Philadelphia, PA, 1993. 24. ASTM E-1570, Standard Practice for Computed Tomographic (CT) Examination, Amer. SOC. Testing Materials, Philadelphia, PA, 1993. 25. Antal, J.J., Dance, W.E., Moravec, J.D. and Carollo, S.F., Experience with an on-off mobile neutron radiography system, Proc. Second World Conf. Neutron Radiog., (eds. J.P. Barton, G. Famy, J.L. Person and H. Rottger), Dordrecht: D Reidel Publishing Co. 1987,407414. 26. Cluzeau, S. and Le Tourneur, P., Stationary DIANE equipment: description and performance of the thermal neutron source, Nuclear Instruments and Methods in Physics Research B, 1994,89,428431. 27. Miller, R.K. and McIntire, P., (eds), Nondestructive Testing Handbook, Vol. 5, Acoustic Emission Testing, Amer. SOC.Nondestructive Testing, Columbus, OH, 1987. 28. Sundaresan, M.J., Henneke, E.G. and Brosey, W.D., Acousto-ultrasonic investigation of filament-wound spherical pressure vessels, Materials Evaluation, 1991,49,601-606,612. 29. Huber, R.D. and Green, R.E., Noncontact acousto-ultrasonics using laser generation and laser interferometric detection, Materials Evaluation, 1991,49, 613-618. 30. McMaster, R.C., McIntire, I? and Meister, M.L., (eds), Nondestructive Testing Handbook, Vol. 4, Electromagnetic Testing, Amer. SOC.Nondestr. Testing, Columbus, OH, 1986. 31. Mahoon, A., Automated eddy current testing of composites, Proc. 20th Intern. SAMPE Tech. Conf., Vol. 20, pp. 517-523, SAMPE, Covina, CA, 1988. 32. Wolfe, W.L. and Zissis, G.J., The Infrared Handbook, Office of Naval Research, Dept. of the Navy, Washington, D.C., 1985.
33. Cohen, J., Elements of thermography for nondestructive testing, NBS Technical Note 1177, National Institute of Standards and Technology, Washington, D.C., (1983). 34. Jones, T. and Berger, H., Thermographic detection of impact damage in graphite-epoxy composites, Materials Evaluation, 1992, 50, 1446-1453. 35. Jones, T.S., Berger, H. and Weaver, E., Large area thermographic inspection of GRP composite marine vessel hulls, Therrnosense XV: Intern. Conf. Thermal Sensing and Imaging Diagnostic Applications, (Lee R. Allen, ed.), SPIE Vol. 1933, pp. 197-206. 36. Jones, T.S. and Lindgren, E.A., Thermographic inspection of marine composite structures, Thermosense XVI: Intern. Conf. Thermal Sensing and Imaging Diagnostic Applications, (John R. Snell, ed.), SPIE Vol. 2245, pp. 173-175, (1994). 37. Vest, C.M., Holographic Inteferomety, New York: John Wiley, 1979. 38. Hung, Y.Y., Shearography: a new optical method for strain measurement and nondestructive testing, Optical Engineering, 1982,21(3), 391-395. 39. Newman, J.W., Shearographic inspection of aircraft structure, Materials Evaluation, 1991, 49(9), 1106-1109. 40. Urabe, K. and Yomoda, S., Non-destructive testing method of fiber orientation and fiber content in FRP using microwave, Prog. Sci. Engng of Composites, Fourth International Conference on Composite Materials, Tokyo, 1982. 41. Bakhtiari, S., Ganchev, S. and Zoughi, R., Microwave swept-frequency optimization for accurate thickness or dielectric property monitoring of conductor-backed composites, Materials Evaluation, 1993, 51(6), 740-743,748. 42. Gray, S., Ganchev, S., Qaddoumi, N., Beauregard, G., Radford, D. and Zoughi, R., Porosity level estimation in polymer composites using microwaves, Materials Evaluation, 1995, 53(3), 404408.
REPAIR ASPECTS OF COMPOSITE AND 39 ADHESIVELY BONDED AIRCRAFT STRUCTURES Anton L. Seidl
39.1 INTRODUCTION
the composites industry, is an unfamiliar language to the uninitiated. Inspectors are often 39.1.1 MANUFACTURABILITY AND at a loss when attempting to describe a condiMAINTAINABILITY OF COMPOSITE AIRCRAFT tion they perceive as a defect; the words STRUCTURE simply do not exist in their standard lexicon. It To the manufacturer, weight reductions, struc- is intuitively clear to even the casual observer tural requirements, manufacturability and that repairs using mechanically fastened conproduction costs have long been obvious pri- ventional materials can be effected quickly, orities. Only recently, however, and only as a under almost any atmospheric conditions, and consequence of persistent user demands, have with minimal investments in tooling, raw maintainability and repairability been added materials, and training. In contrast, repairing to this list. From the operator’s perspective, even relatively minor damage on composite of non-convennevertheless, composite structures continue to structure requires an array tional materials, highly skilled and be a mixed blessing. Clearly, and despite statements being heard to the contrary, the experienced technicians, special tooling and equipment, access to production drawings (to industry would be loath to give up the many locate and interpret the many hidden features obvious advantages gained through the use of characteristic of composite structures), a concomposites and revert to all-metal airplanes. trolled environment in terms of temperature However, the maintenance problems associand humidity, time-consuming preparatory ated with composites cannot be work, cold storage of shelf-life limited and underestimated and may well be regarded as occasionally hazardous materials, lengthy the weak link in the new technology chain. resin cure cycles, post-repair NDT, and legally mandated record-keeping and follow-up 39.1.2 METAL REPAIRS COMPARED WITH activities. COMPOSITE REPAIRS
Compared to the relative simplicity of conven- 39.1.3 COMPOSITE REPAIRS: AN AIRLINE tional metallic structures, composites are PERSPECTIVE replete with complexities that continue to bafThe aim and purpose of this presentation is to fle and confuse maintenance workers trained highlight the principal aspects of composite only in the traditional, i.e. metalworking, structure repairs from an airline perspective. skills. The glossary of terms alone, as used by An attempt will be made to: Handbook of Composites. Edited by S.T. Peters. Published in 1998by Chapman & Hall, London. ISBN 0 412 54020 7
1. describe some of the more common defects and conditions encountered in service;
858 Repair aspects of composite and adhesively bonded aircraff structures subsequent repairs more difficult to perform. When surface defects are detected, it is imperative to determine the extent of the damage that may already have occurred and if possible, evacuate and decontaminate the panel. As 39.2 DAMAGE ASSESSMENT a first line of defense against any further structural deterioration, foil tape should be applied 39.2.1 IMPACT DAMAGE - NON-METALLIC at the earliest opportunity2. STRUCTURE Severe degradation is generally quite obvious, having resulted in visible disbonds and Foreign object impact without skin delaminations. If the damage is due to a recent penetration event, and the process of deterioration has Prior to any repair action, it is important to only begun, the amount of ingested fluid may determine the extent of the damage sustained still be relatively small and if the precise locaby the structure. One must always assume that tion of the contaminant can be determined, the actual damage is more extensive than the complete evacuation and purging may be sucvisible damage'. This is especially true for car- cessfully accomplished. bon fiber-reinforced composites with non-toughened 177°C (350°F) cured matrix Limitations of moisture detectors3 resins. After a foreign object impact, there is generally, but not invariably, some visual indi- Commercial moisture detectors are extremely cation in the form of damaged paint. However, useful devices requiring no special training. because of the elasticity of high modulus Where appropriate, they may be used to deterfibers, the laminate often 'springs back', leav- mine how far any ingested water has spread ing residual subsurface damage in the form of into the core cells adjacent to the point of broken fibers, ply separations and in the case impact. Moisture detectors, however, are effecof sandwich panels, crushed core and dis- tive only on non-metallic (typically glass or bonded face sheets. Tap testing is generally aramid-reinforced structures); they cannot be sufficient to delineate the extent of the damage used on panels containing carbon fibers, or in and should be conducted before removing any zones reinforced with metals. Nor are moisdamaged materials. Defects may also propa- ture detectors effective through surfaces gate during the removal process, often as the coated with carbon-filled conductive paints, on panels having metallic coatings, metalresult of stress relief. coated fibers, or similar lightning protective and EM1 shielding features. Skin penetrations: holes, cracks, tears, gouges, cuts and abrasions
2. give a brief summary of common maintenance practices; 3. describe a limited number of typical repairs.
On the 'wetted' outer surface of the aircraft, even minor penetrations of the face sheet must be regarded as serious because once an opening exists, the part has been rendered permeable to atmospheric moisture and aircraft system fluids. Ingested water can and will degrade the affected part, leading to premature failure. Fluids such as hydraulic oil, when allowed to enter, contaminate both laminates and honeycomb core materials, making
E@cacy of radiography Radiography (X-ray) is presently the only available practical technique for determining moisture contamination in panels containing electrically conductive elements. Given the limitations of moisture detectors and the limited availability of X-ray equipment in the field, water detection by X-ray and subsequent evacuation are generally carried out only during depot level maintenance opportunities.
Damage assessment 859 Interim repair actions - ‘speed tape’ repairs When a composite panel is found to have been penetrated, it is important to prevent further deterioration of the panel. When fluid detrimental to adhesion (hydraulic oils, deicing fluid, engine oils, etc.) is present, the affected area must be thoroughly decontaminated before attempting a repair, or the contaminated material removed entirely. When a permanent repair is to be deferred, fractured material should be trimmed away and the opening covered with foil tape before the aircraft is dispatched to a location where the appropriate repair facilities exist. Foil tapes must be applied with care to prevent their coming loose in flight. Loose foils have been known to create static noises that can interfere with radio communications.
Effects of skin penetration: corrosion, resin plasticization and core dissolution
Any impact damage resulting in skin penetration must be regarded as serious damage. However, unlike non-metallic core materials, which absorb and diffuse water, non-perforated aluminum honeycomb cores tend to keep any ingested water concentrated about the area of the penetration. Left unattended, prolonged exposure will cause the ingested water to migrate to other areas of the panel by gradual, progressive diffusion through the adhesive bondlines and, preferentially, through the core splice adhesives. As the bonding adhesives absorb moisture they become plasticized and their bond strength weakens. At the same time, unprotected areas of the face sheet, doublers, substructural components, cut edges and fastener holes, i.e. where 39.2.2 IMPACT DAMAGE ON METAL-SKINNED the anodic and primer protections have been removed during the manufacturing process, SANDWICH PANELS and machined edges of the honeycomb core, Unlike laminated face sheets, which may are rendered vulnerable to corrosion attack. show little evidence of an impact having Ingested water, if left unevacuated for long taken place, thin metal face sheets (common periods, has been known to initiate chemical on many honeycomb sandwich panels) reactions that lead to complete dissolution of invariably become dented or gouged by for- the aluminum honeycomb core. eign object impact. The resulting surface irregularities are readily seen. Minor damage - no skin penetration
39.2.3 DAMAGED PROTECTIVE COATINGS AND SEALANTS: LEAK PATHS
Shallow dents may be present that do not necessarily result in disbonding of the skin, but there will always occur some crushing of the core cells. A tap test will usually, but not always, determine whether the skin is disbonded. Dents that have not resulted in skin disbonds are generally considered negligible damage and may be filled with an appropriate compound to restore aerodynamic cleanness, provided the added weight does not affect the balance of a critical control surface. Flight control surfaces damaged by hailstones frequently exhibit multiple dents that cannot be repaired by dent fillers without creating an out-of-balance condition.
Water ingestion and fluid contamination must be presumed to exist whenever the protective coatings or sealants of a panel have been disturbed. The cause may be erosion of the protective finish, substrate corrosion, hail damage, minor collisions, or similar foreign object damage episodes. Leak paths, no matter how small, are detrimental to the long-term structural integrity of the panel because they allow atmospheric moisture, aircraft system fluids, or a combination of contaminants, to enter the structure. Subsequent ’ground-air-ground’ and ‘freeze-thaw’ cycling are capable of introducing considerable quantities of water and other
860 Repair aspects of composite and adhesively bonded aircraft structures fluids into the core of a panel. Sandwich panels with thin face sheets of aramid/epoxy are especially vulnerable to moisture contamination via small cracks in the resin gelcoat and at the resin-fiber interface. This phenomenon has been explained as the result of the thermal expansion behavior of the aramid fiber, which is slightly negative in the longitudinal direction and strongly positive in the transverse direction, leading to excessive strain build-up within the weave itself.
Moisture While demonstrably corrosive to metals, moisture is far more pernicious in its effect on composites because it plasticizes resins, degrades their mechanical properties and lowers their glass transition temperature4. The latter effect becomes extremely critical when carrying out hot bonded and/or prepreg repairs that require the heating of the structure. Atmospheric electricity 22-z
39.3 ENVIRONMENTAL DAMAGE AND DEGRADATION
39.3.1 GENERAL EFFECTS OF AGINGz4
All polymeric materials are subject to degradation over time. For this reason, the importance of maintaining protective coatings and sealants cannot be emphasized too strongly. The 'normal' operating environment of an aircraft exposes composite structures not only to considerable static and dynamic loads, but also to significant temperature gradients, extreme variations in humidity conditions, and to a number of chemical agents necessary in aircraft systems, the most detrimental being hydraulic fluid, a powerful solvent. 39.3.2 EXPOSURE OF COMPOSITES TO THE 'NORMAL' FLIGHT ENVIRONMENT
There is abundant evidence that the combined effects of stress, temperature, water and other fluids expose bonded and especially fiberreinforced composite structures to a far wider range of hazards than their baseline metal analogs. Following is a brief description of the most common environmental hazards composite materials are exposed to.
Atmospheric electricity, of negligible consquence to metal structures having inherent conductivities, can have a crippling effect on non-metallics, which compels the operator to place a high priority on periodic testing and proper maintenance of anti-static and lightning protection schemes, i.e. ground paths, bonding fasteners, bus strips, jumper cables, conductive enamel and/or flame-spray coatings, as well as discharge ports. Chemical contamination Aircraft fluids and chemicals that are harmless on metals can effectively destroy a composite, if allowed to penetrate its outer protective layers. Chemical paint strippers routinely used on metal aircraft are occasionally - albeit inadvertently - applied to composite surfaces with destructive consequences, even if the exposure is but of short duration. Presently, composites can only be stripped by abrasive, non-chemical methods5. Overheat conditions Heat, except for annealing temperatures, is of minor concern with metals; by contrast, the heat resistance of compositesis effectively limited by the maximum use temperature of the polymeric matrix. Heat generated by lightning strikes has been known to vaporize matrix resins and create large areas of delamination and fiber fracturing on composite rudders,
Damage removal techniques 861 ailerons, wing and stabilizer tips, nose domes, and nacelle cowling. When exposed to hot gases over long periods, polymeric resin binders, irrespective of chemistry, can become completely destroyed through a process sometimes described as thermo-oxidation. This condition may be found on all types of composites, including those with inorganic matrices, such as metal matrix composites. Preventive maintenance may consist of the application of heat-resistant, ablative or intumescent coatings. Extensive redesign of the detail may be necessary, using metals or, if a fiber composite is to be used, choosing a polyimide or similar high temperature resistant resin system. 39.4 DAMAGE REMOVAL TECHNIQUES
39.4.1 PLANNING THE REPAIR 'THINK BEFORE CUTTING'
After determining the full extent of the damage, the repair technician must consider a range of possible approaches, based on such considerations as damage location, access to the damage, required disassembly to create better access, available tooling and repair materials, as well as the allotted out-of-service time. Because most repairs are 'on-condition', i.e. the result of damage events affecting the structure at unpredictable locations in a multiplicity of manners, allowing only limited pre-planning, the technician's experience and intuitive problem-solving abilities are of paramount importance. 39.4.2 AERODYNAMIC SKIN DAMAGE REMOVAL
If the damage affects the outer, aerodynamic or 'tool' side of a panel and the backside (the %ag side') is accessible, it is best to remove material from the backside, and as much of the core material as necessary, to gain access to the damage. Using this method preserves as much of the aerodynamically 'clean' surface as possible. Repair work is more easily accomplished
from the backside, with the additional benefit of causing only minimal disruption to the aerodynamic surface. Figure 39.1 illustrates this principle. If the backside is inaccessible, the damage must be repaired from the aerodynamic skin side, inevitably enlarging the repair surface and making the repair more difficult to perform. With only one side accessible, the question of how best to apply vacuum pressure is always problematic and requires considerable operator skills. (Applying vacuum pressure for a bonded repair is an art form that must be learned as any other.) As an alternative to field repairs, panels are often removed from the affected structure and routed to a repair facility specially equipped to effect the appropriate restorations. It should be noted that the damages affecting the aerodynamic skin surface normally require 'flush' repairs to preserve the original contour, particularly in zones of the aircraft defined as aerodynamically critical. Except for small damages, the tooling and skill levels required to effect proper repairs do not exist at field stations.
REPAIR PLIES-/
-FILLER
\
REPAIR PLIES-
CORE PLUG
-iti--FLUSH
SIDE
~NON-FLUSH SIDE
Fig. 39.1 Aerodynamic skin side repair.
862 Repair aspects of composite and adhesively bonded aircraft structures 39.4.3 REMOVAL OF METAL FACE SHEETS AND DOUBLERS
Metal-faced sandwich panels are used on wing spoilers, wing and stabilizer panels, flaps, slats, engine cowling, landing gear strut doors, as well as in a multitude of other applications, including aircraft interiors. Damaged, corroded, or disbonded face sheets are generally peeled away after the application of carbon dioxide pellets ('dry ice'). The dry ice is allowed to dwell on the surface until the thermal shock has weakened the bond strength of the adhesive sufficiently to allow the skin to be removed. If done properly, the face sheets separate, leaving the core cells relatively undamaged. 39.4.4 REMOVAL OF COMPOSITE FACE SHEETS AND DOUBLERS
The outer surface of most non-metallic sandwich panels consists of only a small number of prepreg fabric and/or tape plies co-cured onto non-metallic honeycomb core, although precured laminates, secondarily bonded to the core, are also found. Removing damaged or disbonded face sheet materials requires either a rotary sander (physical abrasion) or the use of a hot air gun combined with peeling action. Heating the skin laminate has the effect of weakening the resin fillets sufficiently to allow the technician to peel the face sheet materials with only minimal damage to the core. 39.4.5 REMOVAL OF HONEYCOMB CORE MATERIALS
After the face sheet material has been removed, the condition of the honeycomb core must be determined. Because of the high cost and limited availability of some core materials, repair shops attempt to salvage the original material if at all possible.
Aluminum core removal Severely damaged aluminum core (crushed, corroded, failed node bonds, etc.) should always be replaced. Removal is generally accomplished by using non-metallic scrapers or chisels mounted in a pneumatic rivet gun. Care must be taken to avoid damaging the intact face sheet on the far side. The remaining adhesive fillets on the far side should be abraded with a rotary sander, provided the adhesive is still firmly attached, thils providing a good base for bonding in the replacement core plug. If the far side adhesive is plasticized or unbonded, or corrosion is found between the adhesive layer and the metal skin, the adhesive must be removed for closer inspection and possible reconditioning of the bonding surface. Corroded skins and doublers are routinely replaced.
Non-metallic core removal Non-metallic core materials are generally replaced if crushed or split at the nodes, or if irreversibly contaminated by oil, hydraulic fluids, or other contaminants that would inhibit subsequent repair resin adhesion and cure. Repair shops often attempt to decontaminate core by flushing out the cells with solvent, a method not always successful and a potential environmental hazard. If the contaminant is water, dehydration of the core by evaporation, placing the part in an oven at a low temperature, is often possible, allowing the material to be salvaged (see Section 39.5). When core replacement becomes necessary, the affected sections are generally cut out with knives or rotary cutters; the resin fillets remaining on the far side are then removed with rotary sanders, to create a proper surface for bonding in the replacement core Plug.
Decontamination 863 39.5 DECONTAMINATION
Effects of contaminants on structural integrity
39.5.1 EVACUATION AND DECONTAMINATIONOF POLYMER MATRIX COMPOSITE STRUCTURES
Dimensional swelling and plasticization of the resin matrix generally result from exposure to high humidity at high temperatures, exposure to many aircraft fluids, to chemical paint stripVulnerability of polymers to fluids pers, and to a variety of common solvents. Organic matrix composites typically absorb Absorbed moisture lowers the glass transition between one and two percent of their dry temperature of the laminate*and may be conweight in moisture under normal service con- ducive to additional microcracking within the ditions. There exists a certain risk when such matrix which, in turn, increases the potential assemblies are subjected to the elevated tem- for additional moisture absorption. Microperatures routinely applied during bonding cracks are considered irreversible, since they and laminating repairs. During hot bond remain after the laminate has been completely repairs, the absorbed moisture volatilizes. The dehydrated. Absorbed chemicals may or may effect on the repaired structure may manifest not affect the structural or mechanical properitself in the form of porosities in the bondline or ties of the composite, but generally render the in the laminate. In severe cases, such as when affected part unrepairable because they inhibit water is present in the core cells, the pressure repair resin adhesion and cure. resulting from the entrapped steam often results in uncontrollable skin-to-core disbonding. For these reasons, it is always advisable to 39.5.2 GENERAL PRECAUTIONS pre-dry composite panels when moisture contamination exists in detectable quantities, or Water evacuation under vacuum pressure at may be presumed to exist, given the general elevated temperatures condition of the part. One should keep in mind Removing water is mandatory in all cases, but that fibers, with the exception of aramid, do not the process becomes especially critical if the absorb moisture. Moisture absorption is a phe- repair requires the application of elevated cure nomenon that affects primarily the resin matrix temperatures under vacuum pressure, which and, secondarily, non-metallic core made from is typical or many in situ heating blanket type aramid fibers. As a general rule, resin systems repairs. The operator must be aware that, cured at 170°C (350°F) or above are more resis- under a vacuum bag, lowering the vapor prestant to moisture pick-up than resin systems sure also lowers the boiling point of the water; cured at lower temperatures, which includes all at the same time, increasing the temperature room temperature-curedrepair resins and cold- increases the steam pressure inside the sandbond adhesives-. wich (Fig. 39.2). The result is often a failed repair: blown core and disbonded face sheets. Effect of contaminants on weight and balance Fluids absorbed by, or otherwise introduced into a structure, induce weight gains and may cause out-of-balance conditions in flight control surfaces. Contamination detected should always be evacuated, the leakage paths identified, repaired and the structure resealed.
Removing moisture barriers (coatings and films) Evacuation of composite laminates is best accomplished by first removing any protective coatings and moisture barrier films that may still be present and intact. Barrier materials may be various enamel finishes, sealer
864 Repair aspects of composifeand adhesively bonded aircraft structures 459.02 4 333.89
-
e
,. 145
29.90 20.70 13.98 5.88
1.25
OThrough face sheet evacuation
0.70 0.36
oc OF
-
0.18
0 32
20 68
40 60 104 140
; I
80 176
100 212
120 140 248 284
160 180 320 356
200 392
@zoneof , increasina " risk
160
Fig. 39.2 Pressure and evacuation guidelines for honeycomb core repair.
coats, or bondable plastic films. Their removal is essential to create a path for volatiles to escape rapidly.
Removing face sheet materials
solvents. Often, considerable quantities may be necessary to purge the contaminated core, creating potential environmental hazards. In many cases, complete core replacement may be the only appropriate action.
Evacuation of honeycomb sandwich panels is 39.5.3 SPECIFIC EVACUATION TECHNIQUES most effectively done by removing one of the face sheets. This method exposes the core and Evacuation of fluids from core with face allows the thorough flushing of any contamisheet removed ('open core' evacuation) nants with an appropriate solvent. Complete drying should be performed under vacuum Visible liquids should be evacuated by blowpressure at moderate heat. ing filtered, compressed air across the surface. This should be followed up by flushing the core with an oil-free solvent and then allowing Flushing contaminants with solvents the solvent to evaporate completely. Next, sevEvacuation of chemical contaminants may be eral layers of breather fabric are stacked over accomplished by flushing the core cells with the panel, the assembly is envelope-bagged
Typical repairs 865 and a vacuum of approximately 67 kPa (20 inHg) is applied. The panel is then heated slowly to approximately 74°C (165°F) and allowed to remain at temperature for a minimum of one hour. Evacuation of fluids from core with face sheets intact (‘through-the-facesheet’) evacuation First, all protective coatings and moisture barrier plies must be removed from the areas to be evacuated. Then the gelcoat of the outermost ply should be abraded to expose the fibers. (Fibers inadvertently damaged during this process require subsequent repair.) Next, several layers of breather fabric are applied and the assembly envelope-bagged. Then a vacuum of 34-40 kPa (10-12 in Hg) is applied and the panel heated very slowly (5°C per minute maximum heating rate) to approximately 75°C (165°F) and maintained at that temperature and vacuum pressure for a minimum of 24 h. After this initial drying cycle, the temperature should be increased to 107°C (225°F) and maintained for an additional four hours. Handling of dried details - inspection and storage After drying, details should be re-examined and, if satisfactory, stored in a clean, dry environment until the appropriate repair actions can be taken. 39.6 TYPICAL REPAIRS
the structure. Wet lay-ups normally involve the use of the same type of fabric used in the original construction, in conjunction with a laminating resin capable of room temperature cure under vacuum pressure. The quality of the repair is generally enhanced by applying moderate heat (100°C max) by means of heating blankets, heat lamps or hot air. Heating blankets Heating blankets used in conjunction with vacuum pressure repairs should have an output (watt density) of no less than 7750 W/m2 (5 W/in’). To facilitate draping over curved surfaces, heating blankets with silicone rubber-embedded elements are preferred over mineral fiber-insulated pads, because of their inherent flexibility. Stiffer pads should only be used on flat surfaces. Heat lamps Heat lamps that are used either as the primary heat source, or as a means of augmenting other heat sources, should be 250-300 W tungsten or quartz tube, explosion-proof types. When using heat lamps as the primary source, the effective heat input is controlled by the standoff distance, as shown in Fig. 39.3. To avoid overheating any portion of the assembly being repaired, thermocouples should be placed at several locations to monitor the temperature throughout the cure cycle. The stand-off distance or the positioning of the lamp should be adjusted as necessary to maintain the cure temperature within specified limits.
39.6.1 WET LAY-UP REPAIRS AT AMBIENT OR
ELEVATED TEMPERATURES
Hot air blowers
So-called ’wet lay-up’ repairs are the most frequently recommended because they require only the most basic in terms of equipment, tooling, and repair materials. On the other hand, they are also the most limited in terms of size and applicability because such repairs do not restore the full, pre-damage strength of
Hot air blowers similar to hair dryers are often used to accelerate resin cure; they may also be used for reticulation of unsupported film adhesives. Such devices are typically designed with 1000-2000 W heater elements and fan drive motors.
866 Repair aspects of composife and adhesively bonded aircraft structures
13 __
3. Sand off any protective or decorative finishes and coatings; scrape off sealants, especially silicone sealants. 4 Inspect detail for presence of water or other fluid contamination. 5. Evacuate panel using one of the methods described in Section 39.5. 6. Remove damaged materials - face sheets, doublers, and core, using the appropriate techniques described in Section 39.4.
$ 12 --
Note: Steps (5) and (6) may be inverted, depending on the condition of the part.
14 I 5 l \ n v)
f0
.-
E
v
is
11 --
Core plug repair
I O -_ I
I
I
39.6.2 USE OF ADDITIONAL PLIES OVER WET LAY-UPS
Recognizing that wet lay-ups are inherently inferior to autoclave-cured laminates, many repair specifications9 call for the addition of two or more plies of the type of material used in the original construction, as a means of compensating for the loss of stiffness implicit in wet lay-ups. The added plies do, however, result in weight gain and some loss in aerodynamic cleanness. 39.6.3 TYPICAL WET LAY-UP REPAIR PROCEDURES
Damage assessment and removal 1. Determine perimeter of damaged area by NDTlO. 2. Clean area with solvent.
1. Obtain and prefit replacement core plug, using same as original material, cell size, and density (or an approved substitute). 2. Apply resin compound to edges of core plug to provide a shear tie and insert the plug into the cavity. 3. Apply release film, breather fabric, thermocouples, and vacuum bagging materials. Apply vacuum and check bag for leaks. 4. Cure core splice (shear tie) resin, observing the appropriate time/ temperature relationship specified for the core splicing resin. Maintain vacuum pressure throughout the cure cycle. 5. Remove bagging materials and thermocouples. Face sheet repair 1. Taper and splice joint area. 2. Sand core plug flush with innermost ply. 3. Vacuum up sanding dust, solvent clean repair surfaces and allow solvent to dry completely. 4. Using same as original fiber type and weave style, and observing proper yarn orientation, prepare and impregnate each repair ply of fabric with an appropriate laminating resin mixture. 5. Apply repair plies, observing ply stacking sequence and fiber orientation.
Typical repairs 867 6. Apply perforated release film, breather/ bleeder fabric, thermocouples, and vacuum bag. Apply vacuum and check bag for leaks. 7. Cure laminate under 67-81 kPa (20-24 in Hg) vacuum pressure, while observing the appropriate time/ temperature relationship specified for the repair resin. Maintain vacuum pressure throughout the cure cycle. 8. Remove bagging materials.
9.Oil-free solvents and clean cheesecloth wipers. 10. Heat lamps and/or blankets, hot air gun. 11. Thermocouples and temperature monitoring equipment. 12. Compressed air and vacuum source capable of being regulated. 13. Environmental conditions: Work should be done indoors, under conditions of moderate temperatures (ambient) and low relative humidity (40-65%).
Restoration of coatings, finishes and sealants
39.6.4 WET LAY-UPS USING PRECURED PATCHES
1. Ensure resin is fully cured (must be hard when tapped and resistant to solvents when wiped with solvent-soaked cheesecloth). 2. Reactivate surface by mild abrasion. 3. Clean surface and allow to dry. 4. Reapply finishes, including any anti-static and lightning-protective coatings that may be required. 5 . Reapply any sealants or other coatings removed for repair.
Instead of repairing damaged face sheets 'plyfor-ply', using dry fabrics and laminating resins, prepreg materials may be precured, between layers of peel ply fabric, under autoclave conditions, and stored for later use as patching materials. Precured patches should be perforated to facilitate resin flow and to provide vacuum contact. (Perforations should be of sufficient diameter and spacing to provide a vacuum path and resin bleed, without causing resin starvation at the bondline.) Perforated precured materials may be used with laminating resins, film adhesives, or adhesive paste. When using laminating resins, 3-5 wt% fumed silica ('CAB-O-SIL'TM,made by Cabot Corporation, is generally specified) or an equivalent thickener should be used to improve resin filleting on the honeycomb core.
Materials, tooling, equipment and repair environment 25 1.A two-part epoxy laminating resin of the required chemistry. 2. A compatible core splice resin or compound. 3. Same as original type and style of fabric (unidirectional tape may be replaced with two plies of fabric of equivalent thickness, if allowed by the local structural repair manual). 4. Same as original type, cell size, and density core material, or an approved substitute. 5 . Release film materials, both solid and perforated. 6. Breather and bleeder fabrics. 7. Vacuum bagging film, vacuum gage, and vacuum sealer tape. 8. Abrasive discs, hand-held pneumatic motor.
Limitations of precured materials Precured carbon/epoxy patches are normally applied only over flat surfaces. Precured glass/epoxy patches (or similar low modulus fiber) may be applied over mild curves. Since precured patches are basically 'scab' patches, they should not be used if the surface requires a high degree of aerodynamic cleanness. If precured patches are needed for repairing surfaces having compound shapes, special contour molds must be fabricated and used as a strongback to precure the material, so as to
868 Repair aspects of composite and adhesively bonded aircraft structures produce a precise contour match. The precured patch may be regarded as the composite equivalent of a metal stamping.
39.6.5 PREPREG REPAIRS
Autoclave repairs
Restoring damaged laminates by utilizing the same as original preimpregnated fabric or tape, at the same as original cure temperature and pressure, is normally recommended when full restoration of the original design properAvailable test data indicate that precured ties is a requirement. In practical terms, patches, as well as prepregs and film adhe- however, full restoration should only be sives being co-cured (unless cured in an attempted by depot level facilities, since any autoclave under positive pressure conditions) such action necessitates: should be processed under 3340 kPa (10-12 in Hg) vacuum pressure only. Significant 1. removal of the affected part from its parent assembly; reductions in bond strength have been 2. availability of strongback tooling to mainobserved when such repairs are cured under tain contours; heating blankets and at full vacuum pressure. 3. an autoclave capable of meeting the origiMany combination repair techniques, utilizing nal cure parameters; both precured and resin-impregnated dry fabrics, prepregs and film adhesives have been 4. availability of same as original materials of construction and requisite facilities, equipdeveloped for specific damage conditions and ment and NDT capabilities. damage locations. All such repairs should be cured under reduced vacuum pressure. Clearly, only major operators have the neces-
Necessity of reducing vacuum pressure for bonding precured details and for co-curing prepregs and film adhesives
sary capabilities to conduct what can only be described as a remanufacturing operation. At the present time, only a limited number of the Unlike low temperature (65°C maximum) cur- major airlines have the requisite equipment to ing resins, laminating resins capable of being perform rebuilds to OEM specifications. To cured under vacuum pressure and up to 150°C satisfy market demands, a number of repair (300"F), so-called 'room temperature set/ele- facilities have been granted remanufacturing vated temperature post-cure' resins, produce authority under Part 145 of the Federal high quality repairs and are therefore consid- Aviation Regulations. ered desirable alternatives to prepreg repairs. Because of the hazards inherent in all elevated temperature repairs, especially non-autoclave Non-autoclave repair methods repairs performed under vacuum pressure For limited damage requiring only partial only, a cautious approach is necessary. It is restoration, there are approved alternative almost universally recommended to cure the repair methods. Nearly all these utilize repair under the lowest possible cure temper- prepregs and film adhesives that are normally ature at the expense of elapsed time. A given cured by means of heating blankets under resin may be curable in two hours at 150°C 'vacuum pressure only' conditions. Such (302°F) and may require six hours at 85°C repairs can be carried out with minimal capital (185°F). To reduce the risk of part failure durinvestments. Because such repairs yield lower ing the cure, it is generally advisable to opt for than original design strengths, size limitations the longer cure cycle at the lower temperature. usually apply. These limitations are contingent Wet lay-ups cured at elevated temperatures
upon the specific location of the damage as
Typical repairs
defined in the Structural Repair Manual for the aircraft in question. As a rule, repairs in the vicinity of a load path, as defined by finite element analysis, are severely restricted. The allowable repairs in so-called 'field areas', i.e. at some predetermined distance away from spars, ribs, hinge and latch points, etc., are more generous in terms of size as well as repair method.
Vacuum pressureheating blanket repairs (using prepregs and film adhesives) In situ prepreg repairs are often preferred over wet lay-up/elevated temperature repairs because the resin content of the repair is more easily controlled by using a prepreg. One of the risks associated with the use of production prepregs and adhesive films is that these products were formulated for production and normally require high cure temperatures which, when applied to damaged parts likely to contain residual moisture, may cause severe disbonding of the remaining, thus far undamaged, structure. (The repair action thus severely damages or effectively destroys the
869
part.) It is generally accepted that, before contemplating a prepreg repair, the following factors be given serious consideration. 1. The part must be completely dry (see Section 39.5). 2. If at all possible, the part should be envelope-bagged to prevent backskin disbonding during the cure. 3. The cure should always be effected at the lowest permissible temperature specified for the product. 4. If at all possible, a repair prepreg and/or film adhesive should be selected that is curable at a temperature 40-60"C (104156°F) lower than the original cure temperature. This is of particular importance when repairing structures originally cured in the 170-180°C (338-356°F) temperature range. (Several such products are becoming available as a result of persistent industry demands. Representative products are listed in Table 39.1. This listing is given for reference only and does not imply endorsement of any given product.)
Table 39.1 Repair adhesives and resins curable at reduced temperatures
Product
Class
Manufacturer
Min. cure temperature "C
FM300-2 FM250 FM73 FM123-5
EA9680 EA9394 PL795 CYCOM 919 SP377 Epon9410 DER 329 Epocast35A/927
Film Film Film Film Film Paste Film Resin Resin Resin Resin Resin
Am Cy Am Cy Am Cy Am Cy Hysol Hysol BFG Am Cy 3M Shell Dow Furane
Note: Postcure raises upper use temperature.
"F
120 250 112 230 105 225 95 200 120 250 95 200 120 250 112 235 95 200 80 175 Room temperature R.T. + postcure
Max. use temperature "C
OF
300 180 250 250 300 300 350 160 220
150 80 120 120 150 150 177 70 105 See Note See Note See Note
870 Repair aspects of composite and adhesively bonded aircraft structures
Voids and porosities in vacuum-pressure cured laminates and bondlines Major disadvantages of ’vacuum-pressureonly’ cures are a reduction in the compaction of the laminate and the inevitable formation of porosities in the laminate and/or adhesive bondline. The finished repair yields, as a rule of thumb, approximately only 80% of the strength of an autoclave-cured part in terms of shear and flexural properties. The problem of compacting thick laminates may be overcome to some extent by hot debulking each ply, or a stack of several plies of a laminate, under vacuum pressure before the final cure. This method is labor-intensive but useful; it draws off entrapped gasses, improves resin flow, fiber wet-out and therefore overall laminate quality.
Prepregs co-cured with film adhesives Repair technicians often use a compatible film adhesive together with a prepreg when making a repair. A layer of film adhesive is especially desirable as a bond ply over honeycomb core because it enhances the honeycomb peel strength by providing a deeper glue fillet than would be achieved with prepreg alone. There is, however, beside the added cost, a slight weight gain that must be considered when repairing a weight and/or balance critical part.
Prepregs applied over metal substrates Prepregs applied over metallic substrates always require the use of a layer of film adhesive between the metal and the non-metal. The metallic substrate also requires the normal surface preparations applicable for metal bonding, by one of the methods described in Section 39.6.8. Prepregs are often used to provide debris protection in damage-prone areas of thin-skinned sandwich panels, notably wing flaps and other panelling in line with the landing gear. Occasionally, prepregs are used
as panel edge close-out in preference over metal stampings. 39.6.6 SURFACE PREPARATION FOR NONMETALLIC SUBSTRATES
Abrasion and cutting of plies Taper-sanding is the preferred method of creating a scarf joint at the substrate/repair interface, especially if the substrate material is made from a woven fabric. Repairs in unidirectional tape laminates often use step joints, with each repair ply butted against the original ply. Instead of sanding, the splice joint is then prepared by cutting each ply carefully with a sharp instrument such as an ’Exacto’ knife. It is common to use a lap of 13-19 mm (0.5-0.75in) per ply, although there is lack of agreement with respect to the optimum lap distance or the stacking sequence of fabric plies, i.e. whether the smallest or the largest ply should be placed first. Some authorities calculate the overlap as a function of materials thickness (e.g. L = 187‘) whereas others recommend a straightforward 13 mm (0.5 in) overlap per fabric ply and a 25 mm (1in) overlap per tape ply in the zero degree orientation3,11-13.
Use of peel plies Multi-stage processes using precured laminates often use peel ply fabrics which, upon removal, yield a surface that requires no further cleaning or abrading. Chapter 29 contains some important observations about peel piles. Grit blasting
Grit blasting followed by solvent wiping is sometimes used to prepare non-metallic substrates for subsequent bonding and laminating operations. Plastic media with a Mohs hardness of 3.0-3.5 (US Plastic and Chemical Corporation’s Polyextra and Polyplus granulated plastics, sieve size 30/40,
Typical repairs 871 propelled at a low incident angle (15-30") and at moderate nozzle pressure (25-30 psig) have been demonstrated to remove coatings effectively without damage to fibers, and to leave surface conditions of high quality5. 39.6.7 BASIC REPAIR JOINT PREPARATION
Whatever the specific surface preparation method, the focus must be on producing a smooth, contamination-free, activated bond surface capable of promoting adhesion and, after the cure, capable of transferring the structural loads across the joint with minimal disruption of the load path and minimal stress build-up. Stress risers of any kind, abrupt changes in thickness, brittle adhesives, the wrong scarf angle, poor detail fit-up, preloads, etc. should be avoided. 39.6.8 REPAIR OF METAL BONDMENTS
Honeycomb panels with metal face sheets Because thin-skinned honeycomb sandwich panels are the most easily damaged, structures of this type are most often in need of repair. Several kinds of repair activity are considered typical by the industry: 1. minor repairs consisting of the application of cold or hot bonded metal patches; 2. partial skin and/or core replacement with or without the benefit of autoclave pressure; 3. rebuild or remanufacture (considered depot level repair).
Aluminum surface preparations The quality of the repair is directly related to the quality of the surface to which the adhesive is applied. Poorly or inadequately prepared bonding surfaces are the primary reason why bonded repairs fail.
Cleaning, deoxidizing, anodizing, bonding primer application /cure For optimum joint strength and bond durability, all metal surfaces that are to be adhesively joined require the following essential steps: (1) degreasing; (2) alkaline cleaning; (3) deoxidizing; (4) low voltage anodizing in chromic or phosphoric acid; (5) application, and (6) prebaking of a bonding primer. For other than complete rebuilds, which imply complete teardown of the bonded elements, stripping of all adhesive residues, and full reprocessing of details through solution tanks, tank etching/anodizing and primer prebaking are often omitted at the expense of repair quality and longevity. Comparable values of various prebond surface treatments are shown in Fig. 39.4.
Non-tank anodizing A process known as PANTA (Phosphoric Acid Non-Tank Anodizing) exists but requires extensive preparatory work, equipment and special skills, and has therefore not been fully accepted by the industry at the present time. Parts processed in this manner have been demonstrated to be almost equivalent to tankprocessed parts in terms of bond strength and d~rability'~.
Surface preparation for in situ non-autoclave repairs Typically, the repair patches or partial replacement skins are cleaned, acid etched, anodized, primed and prebaked. Structure not amenable to tank solution processing, i.e. the lap joint areas of the structure being repaired, is typically prepared with an acid paste, followed by a deionized water rinse, air drying, and spray application of a bonding primer without, however, the benefit of prebaking. Elevated temperature prebaking is generally impossible without exposing the structure to heat damage and is therefore
872 Repair aspects of composite and adhesively bonded aircraft structures LOSS IN SHEAR STRENGTH OF 2024 -T3 SAMPLES BONDED WITH FM 123 5 ADHESIVE, AFTER 30 DAYS AT 120OF AND CONDENSING HUMIDITY
-
UNEXPOSED EXPOSED
1 TANK ETCH
I
+ PHOS. ANODRE + CIP ,
I
UNEXPOSED EXPOSED
I TANK ETCH + CIP
OSFD EXPOSED
I
UNEXPOSED EXPOSED
I 0
I 1000
I TANKETCH
ISCOTCHBRITE. MEK + PASA - JEL
I 2000
I 3000
I 4000
I 5000
I
I 8ooo
SHEAR STRENGTH, (PSb
Fig. 39.4 Effect of various aluminium surface treatments on repair bond strength and durability.
omitted, at some sacrifice in terms of bond strength and durabilityI5.
Abrasive cleaning of lap joints For reasons of expediency, many repairs are effected under conditions considered marginal. One common practice is to abrade the joint area with aluminum oxide paper, followed by solvent wiping and the application of the adhesive. Repairs of t h s type, whether the adhesive selected be a paste or a film, are rarely of long duration and should be considered ’interim’ repairs only. On the other hand surface preparations using three-dimensional abrasives such as Scotchbrite@,a product of the 3M Company, in conjunction with high quality bonding
primers, have been demonstrated to produce joints of considerable durability and should be encouraged in preference over abrasion with aluminum oxide paper only.
Application of bonding pressure
Vacuum bagging and bondline thickness control For non-autoclave repairs, the most common method of applying bonding pressure is by means of a vacuum bag. Film adhesives used for repair are normally scrim-supported and thus provide bondline thickness control. (A listing of representative film adhesives available with supporting scrims is provided in Table 39.2. This listing is for reference only and
Typical repairs 873 Table 39.2 Scrim-supportedfilm adhesives Manufacturer
Product designation
Cure temperature range -
"C
O F
-
Max. use temperature "C
O F
250 250 250 200 250
Curing at 120°C (250°F) FM73 FM123-2 FM123-5, FM137 Hysol EA9628 3M AF126-2, AF163-3 AF3109-2 Narmco a Metlbond 1113 Metlbond 1133 B.F. Goodrich Plastilock 7178
107-150 107-120 95-120 113-120 113-120 107-177 95-143 95-135 107-120
225-300 225-250 200-250 235-250 235-250 225-350 200-290 200-275 225-250
120 120 120 95 120
82
180
Curing at 177°C (350°F) Am Cy FM61, FM150-2 FM96 FM300 FM400 FM300-1 FM350 Hysol EA9689 EA9649R 3M AF191 AF131-2 AF143-2 AF147 Narmco a Metlbond 328 Metlbond 329 Metlbond 1515 B.F. Goodrich Plastilock 729-3
163-177 160-177 163-177 163-1 77 150-177 171-182 177-182 175-180 175-1 80 175-180 175-180 175-180 163-19 1 135-185 163-177 171-182
325-350 320-350 325-350 325-350 300-350 340-360 350-360 345-355 345-355 345-355 345-355 345-355 325-375 350-365 325-350 340-360
150 177 177 204 150 177 177 177 177 204 177 150 150 204 150 177
300 350 350 400 300 350 350 350 350 400 350 300 300 400 300 350
Am Cy
See Note See Note See Note
Note: Use temperature increases as a function of cure temperature. a Now marketed by American Cynamid
not an endorsement for any given product.) When using paste adhesives, scrim cloth is normally inserted between the adherends to prevent adhesive squeeze-out and resin starvation in the bondline.
Bondline porosities resultingfiom vacuum pressure The repair technician must be aware that not all film adhesives are equally suitable for bonding under vacuum pressure; indeed, most products are formulated for positive (i.e. autoclave) pressure applications. After curing under vacuum, some adhesives exhibit bond-
line porosities that inevitably result in lowered bond strength, which must be taken into account during the repair design. In an effort to overcome these negative effects, a unique bagging method called 'double-bagging' was developed a number of years ago. This method provides for an inner, 'low vacuum' bag 34 kPa (under 10 in Hg) for expelling volatiles, and an outer, 'high vacuum' bag 81-98 kPa (24-29 in Hg) to provide the equivalent of 4147KPa (12-14 psig) bonding pressure on the assembly. The intent of this method is to minimize the effect of full vacuum pressure on the resin during cure by isolating the laminate within a separate diaphragm.
874 Repair aspects of composite and adhesively bonded aircraft structures
Application of mechanical pressure Mechanical pressure applications are sometimes used when both sides of the part are accessible for clamping. Anacoustical (sound suppression) panels having perforated or otherwise permeable skins make the application of vacuum impossible unless the panel is envelope-bagged. To do so generally requires extensive tear-down of the assembly and removal of the affected panel. Permeable face sheet materials (perforated metal, feltmetal, permeable glass fiber /polyimide laminates, or various fine mesh wire cloth acoustic sheet materials) may be successfully bonded under mechanical pressure, using liquid adhesives, pastes, or unsupported reticulating film adhesives.
Sand bags Pressure application methods employing sand bags, shot bags, etc., exist but are cumbersome and yield repairs of marginal quality and questionable durability. Such repairs should always be rendered 'fail-safe' by the addition of mechanical fasteners to provide a secondary load path should an adhesive failure occur during subsequent flight service.
Pressure damage affects primarily details made from light-weight honeycomb core, which is easily crushed if not adequately protected by well-anchored support blocks. Bondments are especially vulnerable to damage when the integrity of the vacuum bag is breached and compressed gasses enter the assembly. Welldesigned fixtures, proper padding of potential puncture sites, bagging films of high quality, pressure levels appropriate for the materials and part configuration, as well as constant monitoring of the pressure cycle are imperative to prevent damaging parts during the cure. Table 39.3 shows the bonding pressures considered typical. It should be noted that for assemblies incorporating honeycomb (sandwich structures) the recommended pressures are predicated on the compressive strength of a core being simultaneously subjected to both autoclave pressure and elevated temperatures. Panel edges are occasionally collapsed during cure, if not properly supported against side loads. Bevelled edges should have angles between 15 and 20°, as shown in Fig. 39.5. Steeper angles require that the core be stabilized with additional resins or core fillers to prevent collapse under pressure. Application of heat
How to introduce the proper amount of heat for curing the repair adhesive or resin has long A novel pressure application method using an been considered problematic. Unlike manufacinflatable rubber bladder system has report- turing processes, which can be optimized edly been successful16.The loads generated by through cure cycle verification by destructive the inflating bladder must be reacted out testing, most repairs are rather unique and against one or several hard points on the air- influenced by a multitude of factors not easily craft structure, which requires equipment of controlled. Heating blankets of a constant watt model-specific design geometry. density tend to overheat thin sections, e.g. the trailing edges of a bondment, while undercuring the bondlines located over a heat sink, e.g. Specific risks associated with pressure a heavy metal fitting or a spar. It has often Parts may be damaged during repair through been found necessary to protect thin sections improper pressure application. Most damages against overheating by inserting silicone ruboccur in the autoclave and are the result of poor ber pads between the heating blanket and the fixturing and of insufficient attention being part, thus reducing the effective cure temperapaid to proper vacuum bagging techniques. ture in selected areas, while allowing the
lnj7atable bladders
Typical repairs 875 Table 39.3 Recommended bonding pressures
Core material
Thickness/cell size
Density
Max pressure
cm
p.c& kPa in kg/m3 Honeycomb Sandwich and MetaYMetal Panels (AluminumBondments) Aluminum over 12.7 over 0.5 48 or higher 3 or higher 118 84 less than 3 Aluminum under 12.7 under 0.5 48 or higher 3 or higher 35 less than 3 101 No core n/a n/a 169-338 Laminates and Panels Containing Non-metallic Honeycomb Cores Aramid* 3.1 1/8 48-64 3.04.0 118 4.7 3/16 48 or lower 3.0 or lower 84 56-88 635 9.5
No core
3.5-5.5 all
1/4 3/8 n/a
all n/a
Psig 35 25 40 30 50-100
101 84 84 152 or
35 25 30 25 25 45 or
higher
higher
* Nomex HRH, HRP, Hh4X or similar core materials.
colder portions to reach the appropriate cure temperature. Occasionally, heat lamps or other auxiliary means must be employed in conjunc- 3. tion with heating blankets to provide additional heat inputs at critical locations to make sure the resins are fully cured. It is imperative that thermocouples be used at as many locations as necessary to monitor the cure cycle and to ensure the repair meets specification requirements when completed. 4. Specific risks associated with heat
The principal risks associated with repair activities on structures that require the use of thermosetting resins and adhesives are: 1. Water or residual moisture in any portion of the assembly may vaporize and cause additional damage such as ply separation, core node bond separation, or skin-to-core bond failure (see Section 39.5). 2. Overheat conditions may develop under a heating blanket, causing irreversible damage, occasionally a fire. Constant monitoring or the incorporation of overheat
alarms may prevent part damage during the cure cycle. Heat sinks may drain away heat energy required for resin cure, leaving residual uncured materials of unacceptable structural value. Hot bonding should not be carried out during adverse atmospheric conditions or while the aircraft is coldsoaked. Improper heat-up rate control may cause resin flow and gel anomalies resulting in a product of marginal quality. Heat-up rates must be monitored or appropriate control devices used. 5. Foaming adhesives may generate exothermic reactions resulting in irreversible damage. This hazard can be avoided by minimizing the width of splice gaps to be filled through careful sizing and fit-up of details prior to and during lay-up. Nonmetallic core details should be joined by crush splicing rather than by adhesive foams to reduce the amount of reactive polymers present in the panel during the cure cycle.
876 Repair aspects of composite and adhesively bonded aircraff structures
' 0 5052
HRH
15' 80%
100% 100%
86%
30'
45
goo
76%
19%
2%
70%
15%
2%
a
*O
fI O0
b
C
\
0
8'. 1
I
I
15O
30°
-- - - _
450
900
OFF-AXIS BARE COMPRESSIVE STRENGTH OFHONEYCOMB CORE
Effect of core edge bevel on core stability under bonding pressure
e
Fig. 39.5 (a) Bare compressive strength of honeycomb at various angles of loading; (b) Off-axis bare compressive strength of honeycomb core; (c) Effect of core edge bevel on core stability under bonding pressure.
Typical repairs 877 Lower cure temperatures enhance repair safety
be the choice for repairs on aluminum, because of its galvanic compatibility and its favorable Problems associated with hot bonding CTE with respect to aluminum. Graphite is increase exponentially as a function of cure inherently incompatible for the same reasons temperature. Cure temperatures in excess of and may be suitable only for repairing titanium 180°C (360°F) are several times more likely to substrates. Composite repairs can be made by result in a failed part than repairs performed applying multiple layers of prepregged fabrics at lower temperatures. Given the option, or unidirectional tapes, which may be cured by repairs should always be conducted at the means of heating blankets under vacuum or lowest practical cure temperature, using a mechanical pressure. Overlays of this type have suitable adhesive or resin system. See Table been reported to enhance the fatigue life of con39.1 for typical products. ventional metal structures by several orders of magnitudez1.One major disadvantage is the 39.6.9 COMPOSITE REPAIRS APPLIED TO need for a chemically prepared surface involvMETAL STRUCTURES ing the use of acids which, if entrapped under the repair, could cause corrosion and premature Resin-impregnated fiberglass cloth repairs structural failure, making periodic NDT of the on aluminum repair mandatory for the remainder of the airframe life. For maximum effectiveness, Wet lay-up, epoxy-impregnated fiberglass phosphoric acid non-tank anodizing (PANTA) cloth repair patches have been approved repair and the use of bonding primers are essential. methods via OEM Structural Repair Manuals When using graphite as the backbone fiber, a and Military T.0.s for many years1. Utilizing barrier ply of fiberglass is necessary to prevent room temperature curing resins in conjunction galvanic coupling between the repair material with fabrics, wet lay-up repairs can be applied and the substrate. The risks involved have thus over flat as well as curved surfaces with a minfar inhibited the use of graphite on aluminum. imum in equipment and under almost any conditions. Experience has shown, however, that in terms of overall quality and durability, 39.6.10 MECHANICALLY FASTENED REPAIRS these repairs are the least desirable and should be applied only when more advanced methods With increased use of composite materials in are unavailable. All resin-impregnated cloth primary and principal structure not readily repairs over metal require the use of primers removable from the aircraft after a damage (typically nitrile rubber based liquids) to pro- incident, bolted repair concepts are being valimote resin-to-metal adhesion. dated for major skin/stringer and skin/chord damage repairs. Utilizing mainly precured composite elements together with metal douRepairs utilizing advanced fibers blers and splice angles, such repairs can be Repairs utilizing boron/epoxy and graphite/ effected where access is limited to one side of epoxy prepregs over aluminum substrates have the structure only. Essentially, such repairs are been under active consideration for use on mil- a logical extension of, and quite similar to, itary aircraft for some time17-21. Only recently, a conventional mechanically joined metal large freight carrier made the decision to apply repairs, except that both metallic and precured boron/epoxy patches on some of its large trans- composite elements are utilized. The Boeing port category aircraft, malung this the first time B-777 is the first major program to approve that boron is being used for this purpose on a this type of repair on its primary structures, commercial fleet in the USA. Boron appears to chiefly its all-graphite composite empennage. ~"
878 Repair aspects of composite and adhesively bonded aircraft structures 39.7 TECHNICIAN TRAINING AND SKILL REQUIREMENTS
39.8.2 INSUFFICIENCYOF TECHNICAL TRAINING
Personnel engaged in designing and carrying out repairs to bonded and composite aircraft structure should be familiar with the fundamental concepts listed in Table 39.4.
Technical school curricula as well as regulatory guidance materials dealing with composite aircraft repairs are suffering from a technology lag that can, ultimately, only be bridged by greater emphasis on education and training at all levels. Community colleges and vocational schools should encourage the active participation of people experienced in the field (even though they may not possess the requisite academic credentials) and seek to enlist the help of subject matter experts.
39.8 CONCLUSION AND SUMMARY
39.8.1 TECHNOLOGICALEVOLUTION OUTPACING TRADITIONAL AIRCRAFT MAINTENANCE SKILLS AT ALL LEVELS
It has been observed that with every new generation of commercial aircraft, there is an increase in the utilization of composite materi- 39.8.3 DESIGN AIRCRAFT FOR als and a corresponding increase in the MAINTAINABILITY complexity of its design. The transition from simple hand layed-up bonded aluminum hon- Design criteria focused on manufacturability eycomb sandwich and fiberglass-skinned without regard to maintainability may ultiNomex panels to monolithic carbon fiber mately result in compromising flight safety, structures produced largely by means of auto- especially now that composites are finding mated equipment has been a long, inexorable, increasing use in primary and principal airand not altogether painless process from the craft structure. New regulations mandating ’damage-tolerant’ designs should be of great operators’ point of view. value in the determining the design criteria of The end users, principally the world’s future aircraft. Airline customer involvement commercial airlines, are finding it increasin the design of new aircraft must go beyond ingly difficult to keep up with the rapid payload, range, and other marketing concerns; technological changes thrust upon them by the time has come for the designer to solicit the the manufacturers. Despite the large volume comments and suggestions of the maintenance of technical literature available on the subject engineer, the inspector and the mechanic. of composites, there is a dearth of practical information. One is tempted to say that the industry is encumbered by a surfeit of highly 39.8.4 NEED FOR STANDARDIZATION specialized data that is impenetrable to all but the experts. There is no denying that this Standardization of repair methods, practices research is both necessary and beneficial; and especially, repair materials is long overthere is, however, a dire need to make this due. The emphasis must be placed on the data amenable to all through thoughtful dis- typical and generic, rather than the peculiar tillation. The worker in the field must know and proprietary. Cooperative efforts involving the practical effects of this research on the manufacturers, materials suppliers, airlines, daily exercise of his craft. He must be kept repair facilities, regulatory agencies, profesabreast of technological advances and given a sional societies as well as academia will be chance to upgrade his skills in order to meet needed to ensure the long-term viability of composites in aircraft structures. ever-changing demands.
Conclusion and summary 879 Table 39.4 Composite repair training topics Components of Composite Materials Fibers and filaments Glass and quartz; carbon and aramid; boron Fiber; roving; strand; yam; woven fabrics, unidirectional tape; milled fibers Product definitions Fiber mechanical Density; strength; modulus; coefficient of thermal expansion properties Finishes and sizings Chrome; silanes; plasma treatments; resin solutions Specialty fabrics Scrims; peel plies; bleeders and breathers; ceramics Bonding adhesives, Epoxies; polyesters; phenolics; polyimides; bismaleimides; catalysts and hardeners; cyanate esters; acrylics; anaerobics; liquid adhesives; primers; resins, prepregs coupling agents; film adhesives; prepreg fabrics and tape Procedures of Fabrication and Processing Laminating Other structures Curing methods Vacuum bagging Adhesive bonding Sandwich panel construction Core materials
The 'Laminate Code'; isotropic, anisotropic, quasi-isotropic laminates; cross-plied laminates; hybrid laminates; anacoustic laminates Filament/tape winding; RTM; braiding; pultrusions Autoclave; non-autoclave; single and multi-stage cures; postcuring; cure monitoring - flow/gel/set Bagging films; sealant tapes; breathers and bleeders; bagging techniques Pastes; liquids; films; cements; pressure application Face sheets, doublers and close-outs/pans; core; properties of sandwich construction:static strength and rigidity; adhesive filleting; shear ties Metal/non-metal honeycomb; cell sizes and shapes: hexagonal, overexpanded, flexcore; core density/weight; directional properties; compressive and shear strengths Cleaning, etching, anodizing, primer application; aluminum, steel, titanium, other metallic adherends Abrasion; grit blast; taper sanding lap joints; step joints; use of peel plies
Surface preparation for metals Surface preparation for non-steps Joining and fastening Bonded joints; mechanical joints; fastener types and alloys; hole spacing; edge distances; hole sizes for composite joints Sawing; routing; drilling; sanding and grinding; water jet cutting; laser cutting Machining of composites
Protective Coatings and Sealants Pinhole fillers; sanding sealers; primers; surfacers; enamels: epoxy, polyurethane; Organic polymers polysulfide coatings and sealants Carbon-filled enamels; flame spray coatings; EM1 shielding materials Anti-static and lightning protection Silicone coatings; ablative and intumescent coatings; heat-resistant enamels Heat and fire protection Environmental Effects Effect of moisture on uncured resins Moisture Effect of moisture on cured systems Effect of temperature on uncured resins Temperature Effect of temperature on cured systems; glass transition temperature; heat deflection temperature Corrosion Galvanic corrosion: carbon/metal couples Corrosion prevention Effect of rain and particulates Erosion Radiation Ultraviolet; thermal; nuclear
Continued on next page
880 Repair aspects of composite and adhesively bonded aircraft structures Table 39.4 (Continued) Atmospheric electricity Chemicals
Static charges; lightning strikes Aircraft system fluids: oils, hydraulic, deicing Accidental exposures: paint stripper, solvent spills Fatigue and embrittlement effects on composites
Aging Standard Tests for Adhesives and Prepregs Tensile shear; T-peel; honeycomb peel; crack extension (wedge) test Metal adhesion Volatile and resin content; resin flow; gel time; tack Prepregs Interlaminar shear; short beam shear; flexure; 45" in-plane shear; tensile and Cured laminates compressive strength and modulus; sandwich beam; hot/wet strength
Inspection and Quality Controls Visual inspection; tap test; penetrants Non-instrumented X-ray; moisture detector; pulse-echo ultrasonic; through-transmissionultrasonic; Instrumented resonance ultrasonic; eddy current Raw materials handling and storage; environmental controls; processing Process quality controls materials controls; facilities and equipment controls; tool design and alteration; detail preparation; in-process sampling inspections and witness coupons Verification of compliance with specifications:cure cycle chart review; physical Post-repair NDI tests and checks Damage Assessment, Failure Analysis, Preventive Maintenance In-service damage Foreign object impact; environmental degradation Failed repairs 'Lesson learned' Periodic inspections Coatings and sealants Electrical continuity check Specific Repair Methods Repair categories Construction type; original materials; original cure temperature Repair preparation Stripping; damage assessment and removal; decontamination Repair materials Adherends and adhesives; auxiliary/ processing materials selection Surface preparations Metals; non-metals Adhesive/resin cure Applying heat and pressure; cure and postcure; cure monitoring Assembly completion Reassembly; finishes; weight and balance; final inspection Health and Safety Aspects of Composites Chemical exposure Dermal; ocular; inhalation; ingestion routes Hazard levels Acute vs. chronic toxicity Material safety data The MSDS and how to interpret Basic industrial Engineered control systems; personnel protection hygiene Toxicology of Resins and catalysts; solvent and diluents; fibers and fiber dust; sealants and composite materials coatings
References 881 REFERENCES 1. Anon., Advanced composite repair guide, Contract No. F33615-79-3217, Air Force Wright Aeronautical Laboratories, Wright-Patterson AFB, OH 45433,1982. 2. Anon., Environmental Durability of Speed Tape. Summary Report by Boeing Materials Technology, Renton, WA, 1982. 3. Anon., Guidance Material for Design, Maintenance, Inspection and Repair of Thermosetting Epoxy Matrix Composite Aircraft Structures. Montreal/Geneva: IATA, Doc. Gen/3043 1991. 4. McKague, Lee, et al., Test of graphite-fiber sizing effects upon laminate properties. SAMPE I., Nov/Dec, 1979. 5. Egan, William, Composite Paint Stripping Development. Manufacturing Development Report No. 6-35081. Boeing Commercial Airplane Company, Seattle, WA 92124,1985. 6. Hertz, Julius, Moisture effects on the high temperature strength of fiber-reinforced resin composites. Convair Aerospace Division of General Dynamics/Hercules Inc. Joint Study NAS 8-27435’1972. 7. Dexter, H. Benson and Donald J. Baker, Flight Service Environmental Effects on Composite Materials and Structures. NASA Langley Research Center, Hampton, VA 23556. 73rd AGARD Structures and Materials Panel Workhop, San Diego, CA October 7-8,1991. 8. Crossman, F.W. and Flaggs, D.L., Dimensional stability of composite laminates during environmental exposure. Lockheed Palo Alto Research Laboratory, Palo Alto, CA. SAMPE J. July/August, 1979. 9. Anon., Structural Repair Manuals (All Models), Boeing Commercial Airplane Company, Seattle WA, 98124. 10. Seidl, A.L., Inspection of composite structure. Report prepared for ATA/IATA/SAE Commercial Aircraft Composite Repair Committee (CACRC), Washington, D.C. Meeting, December 3-5,1991. 11. Anon., Bonded Component Repair Manual (BCRM), Boeing Commercial Airplane Company, Seattle, WA 98124. Document D651169,1983. 12. Anon., Repair Proceduresfor 250/350 deg. F Cured Aramid Fabric/Epoxy and Aramid/Graphite Fabric/Epoxy Hybrid Composite Structures. Boeing Commercial Airplane Company, Seattle, WA 98124. Document D6-48908,1982.
13. Kuperman, M.H., Graphite/Epoxy Repair Program Test Results, Internal Report for United Airlines, San Francisco, CA 94128,1983. 14. Locke, Melvin C., Non-Tank Phosphoric Acid Anodize Method of Surface Preparation of Aluminum for Repair Bonding. Report prepared for inclusion in Adhesive Bonded Aerospace Structures Standardized Repair Handbook, Technical Report AFML-TR-77-206, Air Force Wright Aeronautical Laboratories, WrightPatterson AFB, OH 45433,1978. 15. Kuperman, M.H., Bond Strength and Bond Durability Study for the Evaluation of Various Surface Treatments for Aluminum Adherends. Internal Report for United Airlines, San Francisco, CA 94128, 1975. 16. Molent, L. et al., Design of an All-Boron/Epoxy Doubler Reinforcement for the F-111C Wing Pivot Fitting: Structural Aspects. In Composite Structures Oxford: Elsevier Science Publishers, 1989. 17. Anon., Adhesive Bonded Aerospace Structures Standardized Repair Handbook Air Force Materials Laboratory (AFSC) ,Wright-Patterson AFB, OH 45433. Technical Report AFML-TR-77206,1978. 18. Baker, A.A., Fibre composite repair of cracked metallic aircraft components - practical and basic aspects. Composites, 1987,18(4). 19. Baker, A.A., Boron Fibre-Reinforced Plastic Patching for Cracked Aircraft Structures. Lecture delivered to the Melbourne, Australia, Branch of the Royal Aeronautical Society. Aircraft, September 1981. 20. Sandow, Forrest A. and Raymond K. Cannon, Composite Repair of Cracked Aluminum Alloy Aircraft Structure. Final Report AD-A190-514, Flight Dynamics Laboratory, Wright-Patterson AFB, OH 45433, September 1987. 21. Kelly, Larry G., Composite Repair of Cracked Aluminum Structure (Fatigue Life Extension Study). Air Force Wright Aeronautical Laboratories, Wright-Patterson AFB, OH 45433, Undated Report 22. Anon., Atmospheric Electricity - Aircraft Interaction. AGARD Lecture Series No. 110. NATO Publication, printed by Technical Editing and Reproduction, Ltd., Harford House, 7-9 Charlotte St., London, UK, 1980. 23. Fisher, Franklin, and Plumer, J. Anderson, Lightning Protection of Aircraft. NASA Reference Publication 1008, National Aeronautics and Space Administration, Washington, DC 20456, 1977.
882 Repair aspecfs of composite and adhesively bonded aircraft structures 24, Springer, George S., (Ed.) Environmental Efects on Composite Materials. Westport, CN: Technomic, 1981. 25. Anon., DoD/NASA Structural Composites Fabrication Guide, prepared under Contract No. F33615-79-C-5125 by Lockheed-Georgia Company for Air Force Wright Aeronautical Laboratories, Wright-Patterson AFB, OH 45433, 1982.
REUSE AND DISPOSAL
40
Harry E. PebZy
40.1 INTRODUCTION
As the United States and the world become more conscious of the shortage of landfill space and the need for conservation, they will be driven to find ways to solve or alleviate the problem by developing methods for reuse or other disposal of materials. Composite materials offer a particular challenge because of the nature of the reinforcements and fillers. In the case of thermoset composites, the matrix or resin cannot be remelted or reprocessed as can be done with thermoplastic matrices. Many of the non-composite thermoplastics are routinely collected and recycled in towns and cities throughout the world. This chapter will primarily deal with thermoset composites because of the challenge of finding methods to put these materials back into use, and in a viable economic scenario. Most of the recycling effort involves reuse of sheet molding compound (SMC) because of its considerable potential in the automotive market and its emerging economic viability. Technology involving SMC is usually applicable to bulk molding compound (BMC)also. Thermoplastic composites will also be discussed. Hybrid composites, presumably because of their relatively small usage and more complex makeup, have had no reported recycling attention, but some of the reuse technologies described herein should be useful. Several practical technologies for recycling and reuse have been
Handbook of Composites.Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
developed but the problem of markets for the reprocessed composites remains largely unsolved. Plastics recycling is generally classified into four categories by final product or by-product, as follows: 1. Primary: Recycling a plastic product into an identical or similar new product; 2. Secondary: Recycling a plastic product into a new product that has less demanding physical or chemical properties; 3. Tertiary: Converting plastic wastes into basic chemicals or fuels through a process such as pyrolysis; 4. Quaternary: Burning the plastics and recovering their heat energy. Composites are recycled in all of these categories, with the fourth, burning for energy recovery, of the least value. 40.2 HISTORICAL PERSPECTIVE
Recycling of certain commodity thermoplastics, particularly polyethylene (PE) and polyethylene terephthalate (PET), has become a routine matter in thousands of cities and towns in the past few years. There have been upturns and downturns in recent years in balancing availability of PE and PET with demand for products and markets. One example is the demand for plastic soda bottles for conversion to polyester for garments (Disenhouse, 1994). The situation illustrates an important point, common to all industries: it is not only a question of developing tech-
884 Reuse and disposal nologies for reuse but making it economically viable to do so. Activity in recycling thermoset composites in Europe, particularly Germany, precedes that in the USA. The action is driven by legislation and public opinion, as is the case in Japan also. 40.3 RAW MATERIAL FORMS
40.3.1 REINFORCEMENTS
Glass fiber Glass fiber is the most widely used reinforcement in the world, with nearly 1.2 million metric tons used in 1992 to reinforce 50 or more different engineering thermoplastic and thermoset resins. Various processes used to recycle or separate the glass from the resin affect the nature and quality of the recovered glass. First quality glass can be reused directly, without further processing, in a new composite material. Other processes yield a glass that will require chemical resizing to renew its interfacial bonding to the resin. When the glass suffers deterioration in strength or contamination in recycling, the glass fiber must be refibered or remelted before it can be used again. Commercially viable technology that will allow widespread use of recycled and contaminated glass in reinforcement manufacturing processes is now being developed. Implementation of this technology will open up glass remelting and fiberization as a viable recycling approach, making it possible to contemplate truly closing the loop for glass reinforcement (Graham,Jutte and Shipp, 1993). In thermoplastic recycling, a mixed stream of resin often occurs. Addition of virgin glass fibers has been shown to provide improved properties in mixed thermoplastics. Use of recycled glass fibers would therefore benefit both thermoplastic recycling as well as glass. The glass could be obtained from recycled thermoplastics or thermosets, either as a regrind or from one of the other approaches.
Carbon fiber Carbon fibers can be recovered from uncured epoxy prepreg scrap by two methods. One is by use of an organic solvent agitated bath which washes out the resin. The second employs thermal degradation, below 400°C (750"F), to eliminate the epoxy. The resulting fibers are cut to length 23 mm (0.10 in) for use with aligned discontinuous fibers and provide improved drapability for complex shapes (Norris, 1990). Because fiber and fiber scraps are randomly oriented, the cutting operation should be carried out in two directions, necessitating a special cutting device (Richter and Brandt, 1987). Aligning short fibers results in composites with mechanical properties that are below those of long fiber reinforced materials, but considerably above those of short fiber reinforced injection and compression molded composites. 40.3.2 RESINS
Work has been done to develop epoxies, polyimides and unsaturated polyesters for recovery (Section 40.4.2). The reversal of chemical reactions utilized to produce thermoset and thermoplastic polymers to regenerate monomers or beginning chemicals is covered below (Section 40.4.4). It is possible to extract polyester resin from uncured SMC and reuse the glass fiber in both SMC and BMC. The extraction is done by solvent extraction with mechanical agitation. 40.3.3 PREIMPREGNATED MATERIALS (PREPREGS)
A multi-year program, beginning in 1991, is underway on carbon/epoxy prepreg scrap recycling at the Center of Excellence for Composites Manufacturing Technology (CECMT), operated for the US Navy by the Great Lakes COmposites Consortium (GLCC). The priority of the scrap program at CECMT is illustrated by the fact that for every
Reuse technologies 885 0.45 kg (1lb) of composite materials used in aircraft, ship or missile manufacture, 0.90 kg (2 lb) of material ends up in landfills. The average ratio is 1.3 to 1. This waste includes both material generated by the material supplier and material generated by the end item manufacturer. In time, composite scrap produced at repair facilities will add to this total. Approximately $1billion of all types of raw prepreg is wasted annually and $25 million is spent on its disposal, according to information prepared for the Naval Industrial Resources Support Activity (Broudy, 1993). Another concern is with the legal status of prepreg waste: is it considered hazardous? Both state and federal regulations must be considered. California, for one, has requirements for cured and uncured waste and its disposal. Other state requirements may vary for both material forms. The objective of the CECMT program is to develop a plan for reclamation of carbon/epoxy prepreg scrap and demonstrate the recyclability of the material, based on a survey of users which revealed that carbon/ epoxy is the major prepreg contributing to the waste stream. This scrap or waste material is sometimes referred to as 'offal'. The aerospace industry generates 60-80% of the prepreg offal, with 50% of the total in broad goods form. The most widely used epoxide is tetraglycidylmethylenedianaline (TGMDA) cured with diaminodiphenylsulfone (DDS). Other accomplishments of the program, as reported by Lause (1993) of GLCC, are as follows. A low speed, high torque shredder was found to be best for cutting high tensile and high shear strength carbon fibers. Other cutting blades evaluated were guillotine, carbide, slitter and shear types. A low torque machine did not give a clean cut. The shredder employs two cutting blades, rotating inwards or counter to one another, powered by a 30 hp motor. Shredder input was the scrap prepreg collected from Gerber cutters, i.e. random size and shape. The offal was cut into 50 mm (2 in)
flakes and molded into test panels. Mechanical and physical properties were found to be 80% of that obtained with virgin prepreg. Work was then begun to use material directly from the shredder, to avoid cutting the high tensile strength material into 50 mm (2 in) flakes. Early indications are that approximately the same strengths can be obtained with the randomly oriented material. Preliminary analysis indicates that most epoxies can be comingled and coprocessed. Niche markets are being sought for uses of the reclaimed prepreg, based on the 80% strength obtained and the comparatively limited amount of carbon/epoxy produced in the US, estimated to be 5.5 x lo6kg (2.5 million lb) annually. Markets are sought which would provide highest profitability for the manufacturer; sporting goods and medical prosthesis devices are possibilities. 40.4 REUSE TECHNOLOGIES
Cured thermoset composites undergo the following steps in preparation for reuse. Cutting complete items ( e g boats) into small pieces suitable for handling for shredding is the first step in the process. Items 1.35-1.80 m2 (15-20 ft2).in size or scrap (cured or uncured) may go directly to shredding. Shredding reduces the material to sizes suitable for the next step, whether it be particle preparation (grinding, milling, granulating) or pyrolysis. Shredding will not be necessary for pyrolysis if the item can be accommodated in the pyrolysis furnace. Classification of the shredded material or of the resulting particles will be required if more than one product is desired (eg. glass fiber and powder). Shredding for most experimental work provides chips (or piecesj50 mm x 400 mm ( 2 in x 8 in) to 50 mm x 50 mm ( 2 in x 2 in); granulation gives 9 mm (3/8 in) and smaller particles; and grinding or milling a powder, suitable for filler, of 50 pm or less. Size of particles and powder is determined by screen size.
886 Reuse and disposal 40.4.1 MECHANICAL GRANULATION OR MILLING
Mechanical granulation is probably receiving the most attention as a method for reuse of the final composite, whether it be waste from the molding process or the cured product. Polyester (unsaturated) SMC, because of its predominant place in the total spectrum of thermoset composites, is the material form of most interest. For example, the total automotive usage for SMC in 1993 is estimated to be 342 x lo6 kg (155.6 x lo6 lb) and the SMC Automotive Alliance projected a 55%jump to more than 594 x lo6 kg (270 x lo6 lb) by 1995 (Wigotsky, 1993).Phenolic and urethane-based thermoset composites have also been demonstrated to be recyclable by mechanical milling to a reinforcing filler. Further, it appears that fully milled thermoset can be reused as filler in an indefinite number of cycles without loss of performance. Mechanical recycling methods are based on using SMC scrap directly without alteration of its chemical nature. The SMC scrap is shredded and/or ground into a form which can be used as a filler. Ball or hammer milling, cryogenic grinding and knife granulating are used for mechanical granulation. A typical granulator uses double cross-angled blades rotating against an opposing angled blade, while a hammer mill reduces material size by the hammer revolving at high speeds and cutting by impact in mid air. The entire SMC article may be ground to a fine powder (including all the glass fiber). The other approach is to shred and mill the SMC article in a manner which will recover some glass fibers which are separated from the particulate material (powder). Recovery of some glass fiber is important from a cost-effectiveness basis because fibers have more commercial value as a reinforcement filler (versus particulate filler). Also, it requires more energy to grind all of the fiber into a powder, then to recover a portion of the glass fibers. Premix found the most economical process to regrind scrap consisted of a
granulator, a hammer mill pulverizer and a classifier. Each piece of machinery is commercially available,: custom design is-not required (Butler, 1991). In the USA, the SMC Automotive Alliance, Owens-Corning and Premix are among the leaders in fiberglass composite recycling, but other proprietary recycling activity is also being carried on (Section 40.10). Canadian work is being done by Phoenix Fiberglass Inc. and Plastiglas Industries (Darrah, 1993). Phoenix Fiberglass, through a mechanical process, is able to separate the two components of a laminate that include the glass reinforcement in the form of fibers and the thermoset resin in the form of a powder. In the pilot plant, the waste laminate was inspected for contaminants before it was introduced into the recycling process, with the result that long and short fibers as well as two types of fillers or resin extenders were recovered. Plastiglas Industries Ltd recycles its own waste composites, but has developed the methodology to combine waste composites with other materials from other sources, as well as with other virgin materials. These other materials may come from the blue box program (a curbside collection effort), from sandblasting operations that want to dispose of their waste from sanding operations, and even crushed concrete. Materials that have been identified for use according to a certain recipe will be sorted by size. Depending on the end-use, pieces that range in size from 6-200 mm (0.254 in) can be used. When the material that is to be used in a recipe is larger than 200 mm (8 in), it is reduced to the required size. This part of the operation may use a specially built watercooled diamond wire saw, which can handle pieces up to and including 3 m (10 f t ) in diameter and/or 3 m (10 ft) square. If further size reduction is required, shredders and grinders are used to achieve the proper particle size A significant German program was sparked by public and political pressure upon
Reuse technologies 887 the plastics industry to become responsible for the reuse and cost of their products at their end-of-life (Schaefer and Plowgian, 1993).The major companies from the SMC/BMC market founded ERCOM Composites Recycling GmbH. The proposed German Refuse Act for car recycling, with its priority for material recycling, also provided impetus for action. ERCOM offers a complete system to close the loop between used parts from automotive service and disassembly plants and the reuse of fibrous reinforcing material in new SMC compounds: 0
0 0
a mobile shredding truck that crushes used parts at disassembly and production sites; transport the compacted material to a centralized fractionizing plant; produce a range of fiber rich recyclate material to sell back to SMC producers and other end users.
Milled SMC is being used in several automotive parts. Up to now there has been no technical development reported of a fiber recovery method. Pyrolysis is another important reuse technology. The Swedish Institute of Composites concentrated on evaluating several proprietary formulations of SMC using two reground fractions: one powdered fraction with sizes in the range of 200 pm - 1mm and one fibrous fraction with sizes of up to 500 pm (Pettersson and Nilsson, 1993).The reground SMC came from automotive bumpers supplied by ERCOM. Existing grinding and fractioning techniques were used. 40.4.2 REVERSIBLE CROSSLINKING
The Department of Energy has sponsored work at Polytechnic University, Brooklyn, New York, to develop experimental thermoset The mobile shredder and fractionizing plant polymers designed for recovery (Tesoro et al., have been in operation since the beginning of 1992). Tesoro has shown that it is feasible to 1992. The mobile shredder reduces the size of recover chemical compounds and polymers large parts to a chip size of about from epoxies, polyimides and unsaturated 50 mm x 50 mm (2 in x 2 in). This is a volume polyesters. Designing the polymers for recyreduction of a factor of five. The commonly cling by introducing labile bonds is utilized for used metallic inserts in automotive parts do epoxy resins crosslinked with disulfide-connot have to be removed beforehand. In fact, taining curing agents and for disulfidepreshredding to this predetermined chip size containing polyimide copolymers. Stressis an essential requirement to assure removal strain curves for epoxy resins cured with a disulfide-curing agent, dithiodianiline of metallic parts in the next processing step. The preshredded material is then processed (DTDA) and a commercial aromatic diamine automatically through a series of pneumatic curing agent (MDA) show the equivalence of and mechanical fractionizing steps. It is essen- the DTDA to MDA. Solubilizationof the cured tial to remove metallic parts in this sequence to thermosets by reduction of disulfide bonds assure a metal free product. A hammer mill is provides sites for re-curing. Epoxies may also used for the grinding step. The resulting fiber- be crosslinked with a polyfunctional reagent rich recyclate can be produced in a series of which reacts with thiol groups formed in fractions containing a mixture of glass fibers, reduction. Tesoro has also worked on obtaining chemical products from cured unsaturated fillers and resin. The emphasis in Japanese composite recy- polyesters (UP) (Tesoro and Wu, 1993). The cling has been in dismantling or cutting up focus of the investigation was on chemical large fabricated parts, particuarly fishing ves- reactions of well established UP structures to sels and boats, bathtubs, and tanks (Kitamura, yield soluble, well-characterized chemical 1993).Shredding is best done by a non-gearing products that would be suitable as raw matetype or a rotary shock shearing type shredder. rials for synthesis. Neutral hydrolysis was
888 Reuse and disposal found to be a viable approach through cleavage of ester bonds, giving phthalic acid in good yield and soluble carboxyl-containing oligomers of potential utility were isolated. Recovered oligomers for acid curing of epoxy resin systems have been explored with encouraging results. Smooth transparent films perhaps suitable for coatings have been obtained. The use of recovered oligomer as saturated acid in the ex novo synthesis of UP resins has also shown promise. 40.4.3 PYROLYSIS
Pyrolysis is the controlled thermal degradation of an organic material into one or more recoverable substances through the application of heat in an oxygen-free environment. Pyrolysis should not be confused with incineration, which is the total oxidation of all the organic materials. Pyrolysis has been used for years to reduce plastic and rubber wastes, particularly the stockpiles of used tires. Pyrolyzing SMC in the USA, however, is relatively new. Primary emphasis to date has been applied to scrap materials in the manufacturing process. Ultimately, recycling of parts at the end of their application life cycle must be addressed if elimination of all SMC from landfill is to be achieved. Pyrolysis has a major advantage in that it is well suited to handling contaminated scrap, i.e. SMC which contains paint, adhesives, fasteners or is mixed with other materials. The SMC Automotive Alliance of the SPI Composites Institute (1991), a group of material suppliers and molders, has worked on pyrolysis of SMC auto scrap. In its early stages, the main problem was shredding and feeding materials into the pyrolysis equipment. A discrete chip with minimum polyester breakup is needed to make the process cost effective. By-product utilization is the other factor which will influence the feasibility of the process for SMC waste disposal. In Japan, The Ship Research Institute (SRI)
carried out experiments in cooperation with Michimae using a batch type external heating pyrolysis furnace (Kitamura, 1993). The system was conducted for scrapped reinforced plastic (RP) ship material. When material is heated for about 10 min, gas is produced and more gas can be collected as the temperature rises. The amount of gas which can be collected reaches a peak at 275°C (527°F).The gas produced is either water cooled or run through a heat exchanger to yield some pyrooil and the rest, gas. The RP used in the experiment had a 40 wt YOof glass content as is common for RP used in ships. The residue is fiberglass and carbon, i.e. 40 wt YO and 12 wt YO, respectively. The Japan Marine Recreation Association under the Ministry of Transport and Sasakawa Foundation has been developing a mobile pyrolysis/incineration system for RP ships since 1990. It is economically advantageous for large structures like RP ships to be treated as they are without pre-processing. The mobile pyrolysis/incineration system comprises two trailers, each of which has a body 7 m (23 ft) long, and 2.6 m (8.6 ft) wide, holding a destructive distillation furnace and a rotary kiln, respectively. The system first pyrolyzes an RP ship into pieces less than 6 m (20 ft) in length under a temperature condition of about 400°C (750°F). After shredding into small pieces, removing metals and moving to the second trailer on which a kiln is loaded, the residue is then incinerated in the rotary kiln. A pilot plant on land has been constructed and tested and mobile pyrolysis/incineration system trailers will be completed in 1993. 40.4.4 CHEMICAL DEGRADATION
Under the proper conditions, the chemical reactions that are utilized to produce thermoset and thermoplastic polymers can be reversed to regenerate the various monomers involved. Among these are the 'chem-olysis' processes: hydrolysis, glycolysis, methanolysis, aminolysis, as well as polyurethane (PUR)
Properties and products of recycled materials 889 foam and reaction injection molding (RIM) depolymerization, saponification, refinery cracking, and thermal 'unzipping' of polyolefins. Since these approaches take the organic polymers back to more basic chemical building blocks, they have the potential advantage of yielding known chemical streams that can be converted, using existing purification and polymerization technology, to new polymer with properties equivalent to the virgin polymer. The depolymerization processes will also free up the glass reinforcement used in the composite. In the polyester field, chemical degradation is being utilized primarily for polyethylene terephthalate (PET).Inasmuch as this thermoplastic is primarily used without reinforcement, e.g. bottles and film packaging, no detailed attention will be given here. Typical products obtained in chemical degradation or depolymerization are dimethyl terephthalate (DMT) and ethylene glycol (EG). Pyrolysis, another form of chemical degradation, has been discussed previously; it is a practical recycling process for unsaturated polyester (SMC).
A disposal/recycling model for composite waste material has been developed by Vital Visions Corp. to help evaluate practical and cost effective disposal or recovery methods (Wood, 1991). The State of Minnesota has funded a study for development of an economic model to determine if collection of fiber reinforced plastic scrap from processors can produce profitable products, either materials or end-use goods (McDermott, 1992).
40.4.5 HEAT CLEANING
40.5 PROPERTIES AND PRODUCTS OF
ical solvent which is blended both in the waste paste and waste resin, to keep them in a pumpable state, Premix has used the waste to fuel cement kilns (Butler, 1991). This waste will be used to fire the kiln and then the ash from the firing is put into the cement. This process also eliminates the need to landfill the ash. Some bathtub manufacturers in Japan use special incinerators to recover the steam for boilers from the waste. However, in this case, secondary treatment equipment is necessary to prevent air pollution. 40.4.7 DISPOSAL/RECYCLING MODELS FOR REUSE TECHNOLOGIES
RECYCLED MATERIALS
Heat cleaning has been utilized in recovering useful products from scrap circuit board laminates (Hanson, 1991). Pennisula Copper Industries, after removing the copper from the board, has found it can heat finished or intermediate thermoset composite products, completely volatilize the epoxy resin, and recover the glass fibers with much of the strength of new fibers. The key is a carefully controlled temperature profile as the scrap moves through the furnace (called a calciner), and control of the process gases. The solidified gases are also potentially useful by-products. 40.4.6 FUEL
Liquid waste from producing SMC and BMC has found use as fuel. By introducing a chem-
Because of the big supply and the big market (automotive), the most effort in thermoset recycling has been in particle recycling of glass fiber reinforced polyester, predominantly SMC. It appears that fully milled thermoset can be reused as a filler in an indefinite number of cycles without loss of performance. Pyrolysis degrades the mechanical properties of the glass fiber but fiber is still useful as a particulate filler. The approaches for the recycling of glass reinforced thermoplastic composites utilize the technology already developed for the non-reinforced polymers used in packaging. However, the thermally induced chemical degradation of thermoplastics and the physical attrition of glass fibers that occurs during
890 Reuse and disposal erties with the regrind. The density decreases, as found by other researchers, in this case about 10% at the highest loading. LORIA@(a registered trademark of Ashland 40.5.1 SHEET AND BULK MOLDING Oil Inc.) and Diffract0 analyses were also COMPOUNDS conducted. The outcome of both tests show Owens-Coming reports in Table 40.1 some that 30 phr material gave the best surface. properties of SMC containing filler made from Beyond the 30 phr loading, the surface deterirecycled SMC (Graham, Jutte and Shipp, 1993). orated. In the BMC evaluation, Premix found very Replacing the calcium carbonate filler with ground SMC actually improves most mechani- little effect on the mechanical properties, with cal properties. Fibers that have been recovered slight decreases in tensile strength, flexural from SMC by mechanical separation show modulus and notched Izod as the regrind loading increased. Owens-Corning has shown even greater potential reinforcement value. Premix Inc. evaluated several formulation that unfractionated, granulated SMC, still conof SMC and BMC, using scrap materials in taining fiber with considerable length and varying ratios (30%, 58% and 8S0/0) (Butler, aspect ratio, provides some reinforcement 1991). Beyond 88%, the material viscosity was value in BMC (Table 40.3). Work by ERCOM in Germany evaluated unmanageable. The SMC formula used for the regrind studies was a typical automotive properties at three recyclate contents obtained formulation (Table 40.2). Their data show that in SMC compounds using SMC recyclate only slight changes occur in mechanical prop- obtained from scrap and used parts (Schaefer repeated extrusion and injection molding require attention.
Table 40.1 Propertics of SMC containing composite filler made from recycled SMC (20% carbon carbonate replaced with composite filler) Standard formulation
Low Density formulation
Control
Recycled
Control
Recycled
Tensile strength, MPa (psi)
77.6 (11250)
97.9 (14 195)
105.6 (15 310)
95.5 (13 850)
Tensile modulus, MPa (psi x lo6)
11 590 (1.68)
12 970 (1.88)
8 927 (1.29)
10 670 (1.53)
Flexural strength, MPa (psi)
186.0 (26 970)
210.0 (30 450)
163.2 (23 665)
195.1 (28 290)
Flexural modulus, MPa (psi x lo6)
8 970 (1.30)
9 867 (1.43)
8 683 (1.26)
9 325 (1.35)
Notched izod, J/m (ft lb/in.)
953 (17.8)
1018 (19)
824 (15.4)
888 (16.6)
Unnotched izod, J/m (ft lb/in.)
1164 (21.8)
1388 (26)
981 (18.4)
958 (17.9)
Water absorption (YO)
0.31
0.28
0.20
0.19
Specific gravity, g/cm3 (lb/fP)
1.76 (110)
1.81 (113)
1.32 (82.4)
1.33 (83)
Properties and products of recycled materials 891 Table 40.2 Typical automotive formulation control PHR
Basted on paste %
BOT %
55 40 5 1.5
1.3 13.69 1.71 0.52 1.37 2.60 59.91 1.37
14.13 10.27 1.28 0.38 1.03 1.95 44.93 1.03 25.00
Polyester resin Polyvinyl acetate LPA Styrene Tertiary butyl perbenzoate Zinc stearate Wetting agent Calcium carbonate Magnesium hydroxide Glass
8 175 4 25%
30 pkr regrind
PHR
Basted on paste %
BOT %
Polyester resin Polyvinyl acetate LPA Styrene Tertiary butyl perbenzoate Zinc stearate Wetting agent Calcium carbonate Regrind Magnesium hydroxide Glass
55 40 5 1.5 4 8 145 30 4 25%
18.83 13.69 1.71 0.51 1.37 2.60 49.91 10.00 1.37
14.13 10.27 1.28 0.38 1.03 1.95 37.43 7.50 1.03 25.00
4
Table 40.3 Properties of BMC reinforced with recycled glass reinforced SMC Tensile
sMc Added
Stress, Mpa
Sample
(%)
BMC control
Elongat ion (%)
(psi)
Modulus, MPa (psi x IO6)
0
27.9 (4 045)
13 099 (1 9)
Stnd. BMC + SMC
6
16.1 (2 334)
Stnd. BMC + SMC
12
BMC Resin +- SMC
70
Flexural
~-
Izod impact
Stress, MPa (psi)
Modulus, MPa (psi x IO6)
Notched,
0.44
96.7 (14 020)
10 548 (1.51)
270 (5.1)
9 858 (1.43)
0.68
68.6 (9 947)
9 789 (1.40)
270 (5.1)
17.3 (2 508)
12 685
0.22
71.4 (10 350)
10 203 (1.46)
209 (3.9)
25.8 (3 740)
7 859 (1.14)
0.36
55.4 (8 033)
6 411 (0.98)
89 (1.6)
(1.W
and Plowgian, 1993). Table 40.4 shows that strength can be maintained and there is a reduction in density. The latter implies a lower weight for a given car part, an added benefit for the auto recycling evaluation underway. In Japan, the RTC (RP Recycling and
J h
cfl lb/in)
Unnotcked,
J h
(ftlb/in)
Treatment Council) of the Japan Reinforced Plastics Society, as the final step in a complete recycling evaluation beginning with cutting up and shredding large RP parts, obtained physical and mechanical property data for SMC where CaCO, was replaced with varying
892 Reuse and disposal Table 40.4 Comparison of properties of SMC compounds using recyclate
Base
5%
10%
15%
Flex-strength, 194 222 185 208 MPa (psi) (28 130) (32 190) (26 825) (30 160) Impact,kJ/m2 102 107 126 145 (ft lb/in) (19.4) (20.4) (24.0) (27.6) Density, g/cm3 1.90 1.85 1.80 1.72 (lb/ ft") (118.6) (115.5) (112.4) (107.4) Note. Recyclate content calculated as a percent of total weight of compound.
amounts of regrind (Kitamura, 1993). Tables 40.5 and 40.6 present these data and confirm results by other investigators that mechanical properties are not harmed (in fact, in some cases improved) and the density is lowered. The Swedish Institute of Composites
(SICOMP) performed mechanical testing on a low-shrink SMC with 20% 25 mm (1in) glass reinforcement and a 50% reduction of the calcium carbonate filler, down to 80 phr (Pettersson and Nilsson, 1993).Improved flexural strength together with slightly reduced flexural modulus was obtained for the SMC containing 10% reground material. Evaluation of a virgin PET polyester and one prepared from reclaimed PET has been carried out by DeMaio (1991).Although the resins did not contain reinforcements and therefore may not truly belong in a book on composites, the test data are of interest while considering SMC and BMC above. Tables 40.7 and 40.8 show that no significant differences of liquid or physical properties exist between the two materials once processed into a polyester resin.
Table 40.5 Physical/mechanical properties of recycled SMC - standard type
Formulation Materials Resin CaCO, Reground GF (1in) (Yo)
100 125 0 30
2 100 78 32 30
3 100 36 60 30
Properties Shrinkage (YO)
0.06
0.07
0.08
Specific gravity, g/cm3 (Ib/fP)
1.73 (108.0)
1.64 (102.4)
1.59 (99.3)
Flexural strength, MPa (psi) Flexural modulus, GPa (psi x lo6)
200 (29 000)
213 (30 885)
186 (26 970)
12.3 (1.78)
11.8 (1.71)
11.4 (1.65)
Tensile strength, MPa (psi) Tensile modulus, GPa (psi x lo6)
78 (11310)
77 (11 165)
79 (11455)
11.8 (1.71)
11.5 (1.67)
11.2 (1.62)
Izod impact, mJ/mm2' (ft lb/in)
82.0 (15.6)
87.4 (16.6)
86.0 (16.4)
*edgewise, unnotched
1
Properties and products of recycled materials 893 Table 40.6 Physical/mechanical properties of recycled SMC - class A SMC
Formulation Materials Resin CaCO, Reground GF (1in), Yo
2 100 180 0 28
2 100 123 38 28
3 100 76 70 28
Properties Shrinkage, Yo
0.00
0.00
0.00
1.85 (115.4)
1.77 (110.5)
1.68 (104.9)
Water absorption, %
0.44
0.39
0.30
Surface quality average deviation (p)"
3.6
3.8
5.6
Flexural strength, MPa (psi)
190 (27 550)
178 (25 810)
167 (24 215)
Flexural modulus, GPa (psi x IO6)
11.5 (1.66)
10.9 (1.58)
10.4 (1.51)
Tensile strength, MPa (psi)
85 (12 325)
82 (11890)
68 (9 860)
Tensile modulus, GPa (psi x lo6)
11.0 (1.60)
10.7 (1.55)
10.2 (1.48)
Izod impact, mJ/mm2' (ft lb/in)
76.5 (14.6)
81.5 (15.5)
78.5 (14.9)
Specific gravity, g/cm3 (lb/ft?
*edgewise, unnotched; "Takeda method.
40.5.2 PHENOLIC (PF) AND POLYURETHANE (PUR) COMPOSITES
to 30% recycled content. The recycled PUR must be ground into a fine powder before reprocessing to obtain the Class-A surface.
While the major activity in thermosets has been with unsaturated polyesters (SMC and BMC), property data have also been reported on recy40.5.3 RECYCLED THERMOPLASTIC cled glass reinforced PF, Table 40.9 (Olson, COMPOSITES 1992)and PUR, Table 40.10 (Graham, Jutte and Shipp, 1993). Work to recycle reinforced PUR Glass-filled polypropylene (PP) has received will increase as structural reaction injection the most attention, perhaps because of its molding (SRIM) finds more use in structural potential for automotive use. Owens-Corning applications and PUR presumably continues to studied the effects of repeated recycling by be the resin most used. Krauss-Maffei granulation of the PP and injection molding (Germany)has a new RIM system that can pro- (Graham, Jutte and Shipp, 1993). Table 40.11 duce exterior auto body panels with a Class-A shows a loss in properties caused by polymer surface from reinforced PUR that includes up degradation and a decreasing fiber length.
894 Reuse and disposal Table 40.7 Mechanical property comparison data;
Table 40.8 Liquid property comparison data
recycled PET
Measured property
Recycled PET
Virgin
Flexural strength, MPa (psi)
117.9 (17 400)
121.4 (17 600)
Flexural modulus, MPa (lo6psi)
3586 (0.52)
3586 (0.52)
Tensile strength, MPa (psi)
65.5 (9 500)
67.6 (9 800)
Tensile elongation, YO
2.5
2.8
Barcol hardness
49
47
HDT, "C (OF)
419 (215)
414 (212)
Izod impact, kJ/m (ft lb/in)
0.16 (3.1)
0.18 (3.4)
Water absorption, OO/
2.0
1.8
PET
Recycled
Virgin
PET
PET
Viscosity, cps
470
420
YO Solids
60.1
60.8
Specific gravity
1.098
1.102
17
20
Gel time, min
6.5
6.0
Cure time, min
1.7
1.5
225 (438)
229 (444)
Measured property
Acid number, solids basis 180°FSPI gel test (catalyzed with 1% BPO)
Peak temperature "C (OF)
glass fiber. The results of additional glass in injection molded recycled PP are contained in Two viable approaches to overcome the prop- Table 40.13. erty loss are the introduction of polymer Jutte and Graham (1991) considered the additives and the use of additional glass. Table effect on properties of PP containing granu40.12 presents data to show the improvement lated SMC. Both coarse and fine recycled SMC in recycled PP with the addition of Polybond were evaluated at three filler levels. Modulus 3001 (BP Chemicals), an acid functionalized was improved and unnotched Izod impact PP. The recycled PP was Himont Pro-Fax 6523 properties decreased; the data are tabulated in with 6 mm (0.25 in) Owens-Corning R-34B Table 40.14. Table 40.9 Mechanical properties of recycled glass reinforced phenolic 40% resin -+ 60% regrind
50% resin + 50% regrind
60% resin + 40% regrind
Regrind fypical properties of virgin material
Density, g/cm3 (lb/fP)
1.45 (90.5)
1.43 (89.3)
1.40 (87.4)
1.61 (100.5)
Compressive strength, MPa (ksi)
252 (36.5)
263 (38.1)
273 (39.6)
230 (33.4)
Flexural strength, MPa (ksi)
100.0 (14.5)
102.0 (14.8)
83 (12)
110 (16)
Notched Izod, kJ/m2 (ft lb/in)
2.05 (0.39)
1.98 (0.38)
2.03 (0.39)
3.6 (0.69)
Unnotched Izod, kJ/m2 (ft lb/in)
7.8 (1.49)
8.6 (1.63)
6.4 (1.22)
8.0 (1.52)
Properties and products of recycled materials 895 Table 40.10 Properties of 10% regrind RIM in glass filled polyurethane
Standard Standard system
< 200pm
Flexural modulus, N/mm2 900 (0.13) (psi x lo6)
836 (0.12)
Tensile strength, N/mm2 (psi)
28.0 (4 060)
24.6 (3 570)
Tensile modulus, N/mm2 (psi x lo6)
750 (0.11)
569 (0.08)
Elongation, YO
160
133
Hardness, Shore D
59
63
1.18 (73.6)
1.20 (74.9)
Density g/cm3 (lb/fP)
+
10% regrind
dation. NBC SRIM systems can be ground, dried and injection molded either neat or blended with nylon-6. Recycling evaluations utilized materials flaked to a granular size of approximately 2 mm (0.08 in). Comparison of NBC SRIM properties to recycled (via injection molding) part properties masks the possible effects of the injection molding process due to the inherent differences in reinforcement length. For this purpose, comparison of unreinforced NBC RIM and injection molded, recycled NBC provides a better understanding of the effect of the recycling process on the resin matrix (Table 40.15). As the test data indicates, the initial injection molding cycle does not adversely affect the properties of the NBC system. When an SRIM composite 3545% fiber reinforcement package was recycled, the recycled SRIM composite yielded properties comparable to virgin injection molded impactmodified nylon, indicated in Table 40.16.
In the course of developing a thermoplastic matrix for SRIM, DSM RIM Nylon Inc. developed a family of nylon block copolymers (NBC) named Nyrim (Reitz, 1992). Nyrim 40.5.4 YIELDS FROM PYROLYSIS chemistry combines AP-caprolactam with an activated ’rubbery’ prepolymer. NBC compos- Because pyrolysis occurs in an oxygen-free ites can be readily recycled by conventional environment, SMC decomposes into three injection molding techniques without degra- recoverable substances: pyro-gas, pyro-oil and
Table 40.11 Mechanical properties of recycled 30% glass reinforced polypropylene ~~~
Sample
~
Tensile
Elongation
Flexural
lzod Impact
Stress, MPa (psi)
Modulus, MPa (psi x IO6)
(”)
Stress, MPa (psi)
Modulus, Notched, Unnotched, MPa J h I/m (psi x IO6) (ftlb/in) (ft W i n )
Glass length, (mm)
Control
80.0 (11600)
6 688 (0.97)
2.19
126.6 (18 355)
7 562 (1.10)
103 (1.93)
421 (7.90)
361
Recycled
69.6 (10 090)
6 329 (0.92)
2.28
106.4 (15 430)
6 749 (0.98)
80 (1.49)
321 (6.01)
329
IX
Recycled 2x
63.5 (9210)
6 026 (0.87)
2.29
100.0 (14 500)
6 306 (0.91)
74 (1.38)
292 (5.47)
305
Recycled 3x
55.9 (8 105)
5 461 (0.79)
2.47
89.8 (13 020)
5 557 (0.81)
64 (1.20)
253 (4.73)
280
Recycled 4x
52.0 (7 540)
5 261 (0.76)
2.50
86.1 (12 485)
5 288 (0.77)
60 (1.12)
241 (4.51)
271
896 Reuse and disposal Table 40.12 Propertics of polybond modified 30% recyclcd glass reinforced polypropylene
Sample
Control Recycled l x
+ 5% Polybond
+ 10% Polybond + 15% Polybond
Tensile
Elongation
Stress, MPa (psi)
Modulus, MPa (psi x IO6)
'%)
76.5 (11090) 62.5 (9 062) 73.1 (10 660) 75.8 (11000) 77.2 (11200)
6 405 (0.93) 5 798 (0.84) 6 143 (0.89) 6 260 (0.91) 6 316 (0.92)
2.31 2.45 3.64 3.69 3.45
Flexural
lzod impact
Stress, MPa (psi)
Modulus, MPa (psi x 109
Notched,
(ftIb/in)
cft lb/in)
121.7 (17 650) 97.6 (14 150) 116.1 (16 830) 120.9 (17 530) 124.2 (18 000)
7 239 (1.06) 4 888 (0.71) 5 681 (0.82) 5 632 (0.81) 5 729 (0.83)
101 (1.89) 73 (1.36) 112 (2.10) 121 (2.26) 126 (2.36)
400 (7.49) 314 (5.88) 595 (11.14) 638 (11.94) 679 (12.72)
J h
Unnotched,
J h
Table 40.13 Properties of increased glass content recycled reinforced polypropylene
Sample
Tensile
Elongation
Stress, MPa (psi)
Modulus, MPa (psi x 104,
(%)
Control (30% glass)
76.5 (11090)
6 405 (0.93)
Recycled l x (30% glass)
62.5 (9 062)
+ 2.5% glass
Flexural
lzod impact
Stress, MPa (psi)
Modulus, MPa (psi x IO6)
Notched, cft lb/in)
cft lb/in)
2.31
121.7 (17 645)
7239 (1.05)
101 (1.89)
400 (7.49)
5 798 (0.84)
2.45
97.6 (14 150)
4 888 (0.71)
73 (1.36)
314 (5.88)
67.2 (9 744)
6 260 (0.91)
2.36
101.5 (14 715)
5460 (0.79)
81 (1.52)
317 (5.94)
+ 5.0% glass
70.3 (10 200)
6 964 (1.01)
2.13
108.8 (15 775)
6101 (0.88)
82 (1.52)
302 (5.66)
+ 7.5% glass
73.8 (10 701)
7377 (1.07)
2.08
112.2 (16 270)
6722 (0.97)
85 (1.59)
305 (5.71)
+ 100% glass
75.8 (11000)
8 067 (1.17)
1.87
117.4 (17 020)
8 169 (1.18)
83 (1.55)
281 (5.26)
solid by-product (essentially glass fiber and filler). In one test by the SMC Automotive Alliance, the resulting composition was gas 14%, oil, 14% and inert solid 72%. The gas generated was sufficientto fuel the pyrolysis unit, making it self-sustaining. The energy content was close to natural gas. The solid by-products can be processed into fillers for SMC, BMC or thermoplastics. When milled into a filler, the
I/m
Unnotched,
I h
pyrolyzed SMC can be recycled successfully into general purpose and Class A automotive SMC, at up to 30% of the calcium carbonate filler loading with no adverse effects on processing or physical properties (Rusch, 1993). The pyrolysis experiments performed by Michimae in Japan on scrapped RP ship material gave the yields in Table 40.17. The pyro-oil which was recovered had a strong acid content
Properties and products of recycled materials 897 Table 40.14 Propertics of polypropylene containing granulated SMC
Sample
Tensile SMC
0
Polypro pylene + coarse recycled SMC
15 30 50
Polypropylene + fine recycled SMC
15 30 50
Polypro0 pylene + 9% chopped glass
Flexural Stress, MPa (psi)
Izod impacf
Modulus, Notched, Unnofched, HDTUL MPa l/m 'Im "CPF) (psi x IO6) Cft lb/in) cft lb/in)
Stress, MPa (psi)
Modulus, MPa (psi x IO6)
22.13 (3210)
1035 (0.15)
9.72
29.17 (4230)
965 (0.14)
77.9 (0.14)
968.5 (18.14)
63.3 (146)
21.56 (3 120)
1724 (0.25)
3.50
33.51 (4 860)
1310 (0.19)
84.4 (1.58)
236.5 (4.43)
67.2 (153)
21.03 (3050)
2 207 (0.32)
2.34
34.83 (5 050)
1655 (0.24)
89.2 (1.67)
155.4 (2.91)
78.3 (173)
22.43 (3250)
3 034 (0.44)
1.61
36.21 (5250)
2276 (0.33)
101.4 (1.90)
148.4 (2.78)
94.4 (202)
20.55 (2980)
1448 (0.21)
4.78
32.41 (4 700)
1310 0.19
74.2 1.39
205.0 (3.84)
51.9 (151)
20.96 (3 040)
2 069 (0.30)
2.47
33.93 (4 920)
1.655 (0.24)
82.2 (1.54)
163.9 (3 07)
84.4 (184)
17.17 (2 490)
2 828 (0.41)
1.25
3.076 (4460)
2483 (0.36)
73.14 (1.37)
104.1 (1.95)
82.7 (181)
37.93 (5500)
2 000 (0.29)
3.45
50.69 (7350)
1448 (0.21)
150.6 (2.82)
361.4 (6.77)
114.9 (239)
(%)
Polypropylene
Elongation (")
Table 40.15 Properties of recycled neat NBC RIM
Table 40.16 Properties of recycled NBC SRIM part
Property
Virgin
Recycled
Tensile modulus, MPa (psi x lo6)
2 445 (0.35)
2 480 (0.36)
Tensile modulus, MPa (psi x lo6)
10 400 (1.51)
Tensile stress, MPa (psi)
58 (8 410)
60 (8 700)
Tensile strength, MPa (psi)
120 (17 400)
40
125
Flexural modulus, MPa (psi x lo6)
2 485 (0.36)
2 450 (0.35)
Izod, kJ/mZ (ft lb/in)
10
(1.9) 2
10 (1.9) 1.2
Melting point, "C ("F)
214 (417)
220 (428)
Density, g/m3 (lb/fP)
1.10 (68.7)
1.14 (71.2)
3.1
3.1
Elongation, YO
Mold shrinkage, YO
Water absorption, YO
Elongation, YO
3
Flexural modulus, MPa (psi x lo6)
8 000 (1.24)
Notched Izod, kJ/m2 (ft lb/in)
10 (1.9)
Values are dry, as molded
with a viscosity which resembled heavy naptha, a flash point close to that of gasoline and was highly flammable. The Government Industrial Research Institute and Shikoku/MITI conducted a pyrolysis experiment using steam in a gas-heated continuous-batch method, with temperatures
898 Reuse and disposal of typically 500°C (932°F). RP from ships containing 40% fiberglass was used. The system produces hardly any smoke or gas. The resin yields pyro-oil and fiberglass can be recovered without loss of strength. Table 40.18 shows the yield (Kitamura, 1993).
Table 40.17 Yield from pyrolysis - SRI/Michmae Materials
Yield
wt.%
Gas (liters) Pyro-oil (g) Remainder (8)
90.7 344 524
34.4 52.4
Note: per 1000 g of RP
40.5.5 HEAT CLEANED REINFORCEMENTS
Peninsula Copper Industries (PCI) has obtained data on their heat-cleaned glass fiber incorporated into polyester, polypropylene and nylon matrix laminates (Hanson, 1991). The performance of recovered glass fiber reinforcement is comparable to virgin glass in many respects in polyester and nylon. Table 40.19 compares PCI recovered fiber with PPG fiber with a polyester resin. Table 40.20 presents polypropylene data, comparing the base resin properties to those using 22 wt YO bare reclaimed glass and two formulations of 22 wt % silane-treated glass. The difference between Formula A and Formula B is not explained, but
Table 40.19 Average mechanical properties of polyester /PCI-Glass composites and PPG composite.
Table 40.18 Yield from pyrolysis - GIRIS/MITI
wt.%
Materials Cas (CO) Pyro-oil: Styrene monomer Solids (phthalic) Remainder: Fiberglass Carbon
Property
18
Tensile strength, MPa (psi) Compressive strength, MPa (psi) Flexural strength, MPa (psi) Flber length, cm (in)
25 11 40 6
Note: per 1000 g of RP
PCI
PPG
25.0 (3 626) 89.7 (13 000) 81.4 (11 800) 10.2
24.9 (3 600)
(0.4)
Table 40.20 Average mechanical properties of polypropylene/PCIglass composites Property Unfilled PP
Filled, untreated
Filled, treated A/B
Tensile strength, MPa (psi) Tensile modulus, MPa (psi x lo6)
24.70 (3 582)
26.68 (3 869)
27.20/31.67 (3 945/4 593)
873 (0.127)
2 120 0.308)
2 540/1370 (0.369/0.199)
Yield stress at break, MPa
25 (3 630)
20.1 (2 910)
(2 602/4 121)
Strain at break (YO,at 50 mm/min)
556
391
313/16.7
Glass content (volume YO)
0.0
22.0
22.0
(psi)
17.9/28.4
n/a n/a 84.5 (12 300) 6.4 (0.25)
Disposal of nonrecyclables 899 Table 40.21 Average mechanical properties of nylon 6/6 PCI-glass composites
Property
WPCI
glass fiber Tensile strength, MPa (psi)
102.8 (14 905)
Flexural strength, MPa (Psi)
184.2 (26 690)
Flexural modulus, MPa (psi x lo6)
5292 (0.77)
b o d impact strength (notched 1/8 in), kJ/m (ft lb/in) (unnotched), kJ/ m (ft Ib/in)
(0.8)
0.36 (6.7)
Elongation, YO
3.0
Mold shrinkage, (in/in, Yo)
0.04
0.006, 0.6%
Glass content (vol Yo)
14.8
both show improvement over the base resin and the bare reclaimed glass. The values overall fall far short of most of the commercially available glass filled PP, where tensile strengths range from 29 MPa (4200 psi) to 82.5 MPa (12 000 psi) with a mean value of 44.7 MPa (6500 psi.). The performance of reclaimed silane-treated glass fibers in nylon is closer to that obtained for commercial compounds than for polypropylene. Properties obtained using PCI glass are given in Table 40.21. Depending on the resin formulation, manufacturer’s data show tensile strengths for virgin glass reinforced nylon range between 75 MPa (11000 psi) and 151.8 MPa (22 000 psi). Flexural properties show a similar relationship. 40.6 DISPOSAL OF NONRECYCLABLES
40.6.1 INCINERATION AND LANDFILL
The more desirable forms of disposing of composites are described in the previous section, both for economic payback and avoiding
adding to landfill. Incineration leaves a residue which normally is deposited in a landfill, but increasing legislature action as well as public opinion is reducing the use of landfills. The energy content in SMC can be recovered by incineration, but the relatively low energy content together with the high amount of ash residue from the incineration makes incineration of SMC quite unprofitable. The ash residues from incineration are not useful as fillers in new SMC since the high temperatures when incinerating, together with the presence of oxygen, converts the calcium carbonate to calcium oxide and that will adversely influence the maturation process (chemically thickening) when making new SMC. There are no disadvantages, however, either in the form of added pollution or toxic emissions. The important aspect that SMC consists of only 20-30% organic materials makes incineration and chemical degradation less suitable compared to particle recycling (regrinding) and pyrolysis which are using the inorganic part of the material as well. The non-metal mixture that remains, known as ’fluff’, after such items as automobiles and white goods (refrigerators, washing machines, etc.) are processed in shredders, is normally disposed of in landfills. Seventy-five percent of shredder plant feed is from cars. This feed mixture has a composition of approximately 15-25% plastics, 25-75% inerts and 2-35% moisture. Research by General Motors, Ford and Chrysler and the American Plastics Council is being conducted on recovery, recycling and reuse of plastics from scrapped cars and trucks. Although plastics residue from recycling vehicles represents less than 1%of the solid waste sent to landfills in the USA each year, auto industry officialsestimate that it still accounts for more than 1 million tons of the nation’s solid landfill waste annually. Research efforts are focused on materials that currently are not recovered, primarily plastics, glass, fluids, sealers, fabric, adhesives, paint, and rubber (fluff). Fluff accounts for less than
900 Reuse and disposal 2 wt YO of municipal solid waste. The joint effort targets the plastics portion of the fluff. The recovery of plastics from shredder residue is hampered by frequent contamination from paint, other plastics, metals, and adhesives. Thus another goal of the effort is to develop better methods to purify scrapped plastics. The Department of Energy’s Argonne National Laboratory has developed a process to separate plastics and other recyclable materials from the ’fluff’ of shredded automobiles. The process utilizes a fluidized bed process and yields PUR, fines, iron oxide and mixed plastics. Germany appears to be moving away from its rules forbidding incineration of many kinds of waste. The Environment Ministry reportedly will introduce legislation that would allow incineration of composites and certain plastics if recycling proves too expensive (Protzman, 1993). 40.7 DESIGN FOR REUSE 40.7.1 AUTOMOBILES
An obvious approach to assist in the economic success of recycling of multi-material parts is to design the item originally for ease of disassembly. Of particular interest are automobiles and white goods (e.g. refrigerators). The major auto makers of the US Council for Automotive Research (USCAR), through their Vehicle Recycling Partnership, are studying ways to enhance compatibility and dismantling of plastic parts (Wigotsky, 1993). Some bumper parts now need only four bolts instead of twelve. Potential amendments to the US Resource Conservation and Recovery Act, passed in 1976, may specifically name automobile solid waste. Federal regulations could possibly require certain car design features, if only by excluding certain materials. The VRP is working to establish a strong infrastructure for recycling plastic parts, starting with things that are easy to get at, such as fender liners, fan shrouds and radiator supports. A well-estab-
lished infrastructure for recycling could head off severe government restrictions on plastics use. As with durables in general, aids for recycling vehicles include design for easy separation and collection, parts consolidation to simplify separation, material identification on parts, restriction of plastic parts to recyclable resins, and avoidance of paints, coatings and heat sensitive additives. Automakers worldwide have agreed on an identification code for all plastic parts heavier than 100 g (3.5 oz). Ford Motor Company has issued guidelines on design for for recyclability (Miller, 1993) Germany has been a leader in design for reuse. Porsche, on assignment from the Automotive Engineering Research Association, investigated designs for mediumpriced autos in which plastic parts could be readily dismantled and recycled. The best designs for meeting these objectives were those in which combinations of easily dismantled parts were manufactured from a single recyclable resin. An example of such a combination is a PP bumper system with a PP sheath, an expanded PP core and glass-matreinforced PP beam (Nir, Miltz and Ram, 1993).Many reclamation projects focus on the bumper system. Bumpers are collected in pilot programs, ground into particles and blended with virgin resins. Use of a single recyclable resin has been practiced by Peugeot, Ope1 (single resin bumpers) and BMW (body panels). The use of one resin in different forms, as described above, is also under development by Reko and DSM in auto dashboards. In Landshut, Germany, BMW has a facility capable of dismantling 25 cars a day. Eighty percent of the parts in the BMWs low end ‘3 series’ can be reutilized in some form. Plastic bumpers are turned into interior wall panels. Seat covers and PU foam are used as sounddeadening material for the floor in the ’3 series’. BMW also has three dismantling plants in the USA (New York City, Los Angeles and Orlando) (Protzman, 1993).
Applications and markets 901 40.7.2 APPLIANCES
roof-rack module which uses recycled filler. Introduced in 1992, the rack may have been Design for recyclability, of appliances and the first use of recycled filler in new SMC. other white goods, is underway at several Germany’s ERCOM and the SMCAA have companies. Such designs allow for easy disassigned an agreement to ’enhance the use of sembly, e.g. snap fit assembly with no SMC/BMC in automotive parts by providing fasteners, with each part preferably manufacan organization in each country that will tured from a single, readily recycled resin and accept scrap parts for recycling’. SMC is well labeled as to its composition. Polymer ahead of thermoplastics in automotive recySolutions, a joint venture of GE Plastics and cling use for major parts. This is because SMC Richardson Smith, has already designed an fillers appear to be continually reusable witheasy-to-disassemble refrigerator and a circuit out sacrificing physical properties; replacing breaker box (Nir, Miltz and Ram, 1993). half the CaCO, with SMC filler reduces part weight and cuts compound cost. Thermoplastics have limitations on the 40.8 APPLICATIONS AND MARKETS amounts or frequency of reprocessing permitted before significant mechanical property 40.8.1 AUTOMOTIVE degradation occurs. There is also concern The automotive industry is evaluating many with possible mixing of incompatible resins. applications of recycled composite materials, Plastic material suppliers are working to proparticularly SMC, in new auto parts. The tech- duce resins with recycled plastic content nology to recycle SMC, as discussed earlier, is (Miller, 1993). available and apparently viable. The major There has, however, been reported one barrier to recycle SMC is the lack of an infra- successful application of a recycled thermostructure, i.e. collection, processing, and plastic composite. Chrysler will feature distribution, but this network will develop recyclable fenders for all of its 1993 LH pasbecause of environmental pressures, etc. senger cars. The fenders are made of The first production application of recy- DuPont’s Bexloy K, a new glass-reinforced cled SMC filler in the USA was by General PET polyester based composite. They can Motors on the 1993 Corvette, an inner panel either be reground and molded again or for the rear panel assembly (Rusch, 1993).The returned to the two pure feedstock ingredipart does not require a Class A surface. ents through a patented DuPont Corvette anticipates expansion of recycled methanolysis process (American Plastics SMC to eight more underskin panels in ensu- Council, 1993). ing models. General Motors spokesmen predict that In other applications, the SMC Automotive within five years there will be the necessary Alliance (SMCAA) reported Ford Motor Co. infrastructure for recycling such parts as expected to start using SMC with recycled fender liners, HVAC ducts, fan shrouds, radiacomposite filler on the Econoline Van interior tor supports and many interior components engine cover during the 1993 model year, and (Wigotsky, 1993). Chrysler will use recycled SMC on interior trim on the RAM Van, the first parts that are 40.8.2 BUILDING CONSTRUCTION visible. And in Europe, the SMCAA listed three production auto parts that contain recy- A Canadian company has developed a prodcled SMC as filler (Audi spare-tire well, VW uct Stratum (Stratum@ is a registered Polo engine module and VW Passat front-end trademark for the recycled materials produced panel). Japan’s Toyota Carib has an interior by Plastiglas Industries) composed of a plastic
902 Reuse and disposal binder, usually polyester, filled with up to 85% by volume of recycled composite scrap (Darrah, 1993). One application encapsulates the scrap-resin mixture between two FRP skins. The resulting panel has passed the full scale fire test for exterior panel systems. Work in Japan has been done to develop a gypsum filled with milled composite thermoset scrap and a mortar with milled composite scrap, both with mechanical properties suitable for construction (Kitamura, 1993). 40.8.3 ELECTRICAL PARTS
In Germany, two parts made with new SMC formulations containing 10-15% by weight fiber rich recyclate are already in production or approved for production. These are an electrical distribution cabinet and a cable distribution base (Schaefer and Plowgian, 1993). Strength and modulus can be maintained with a reduction in density. 40.8.4 OTHER APPLICATIONS
Peninsula Copper Industries has identified an application for small chips of printed circuit board in an epoxy/coal tar base (Hanson, 1991). This composite is used as a cushion material between rails and the steel approach plates at railroad crossings. Initial tests indicate that this composite is an effective replacement for rubber-based materials presently used. Several parts from SMC, made from recycled SMC, have gone into production or are approved for production in Germany. These include sinks, seats and chairs. In the thermoplastic field, a commercial application for ‘Nyrim’ nylon block copolymer-based SRIM is a manhole cover (Reitz, 1992). It utilizes a 3.545% fiber reinforcement package. The recycled SRIM composite yielded properties comparable to virgin injection molded impact-modified nylon.
40.9 LEGAL AND ENVIRONMENTAL
ASPECTS
The major federal waste legislation in the USA addressing solid waste issues is the Resource Conservation and Recovery Act of 1976. It requires the Environmental Protection Agency (EPA) to set guidelines for government procurement of recycled products and it mandates Federal agencies and contractors to implement affirmative procurement programs. Most states and many local governments have established programs to procure recycled materials. The 1976 Act was reauthorised in 1988 and continues in effect (Donnelly, 1993). An industry concern is that the federal regulations might venture into the areas of car design, if only by excluding, within certain time periods, certain materials and design approaches. In the meantime, Senator Baucus (D-MT), Chairman of the Senate Environment-Public Works Committee, unveiled in April, 1993, a four-part blueprint for recycling to be part of a recycling bill to be introduced later. Two parts concern composites: the Federal Government should take the lead in procuring goods made from recycled materials and second, manufacturers should be responsible for their products when they become waste. In Germany, increasing concern for the environment has led to proposed laws for the reuse of post consumer and post industrial waste. Legislation for recycling of consumer goods packaging was enacted in 1992, setting minimums for collection and recycling through 1995. Now proposed is the German Refuse Act for Car Recycling, with its priority for material recycling (Schaefer and Plowgian, 1993). Other proposed legislation would levy a tax on new cars to pay for their eventual disposal. Another proposal by Germany’s Environmental Minister would make automobile manufacturers responsible for the final disposal of their cars. Still another stipulated that by 1993, plastics in new cars produced in
Organizations active in composites recycling 903 Germany would be required to contain 25% by weight recycled materials. Legislators and industry will be following Germany’s new environmental laws, expecting to learn from that country’s experience, especially since the current Democratic administration has expressed interest in a progressive environmental policy, i.e. one of reduce, reuse and recycle. Two new laws have been enacted in Japan in 1991; one is ’Law Concerning the Utilization of Recycled Resources’ and the second ’Amendment of Waste Processing and Cleaning’. Earlier, the National Diet amended a part of the Marine Safety Act to address the problem of treatment for scrapped ships. Research and development on dismantling, cutting up, shredding and various recycling methods of other large items followed. Such items include bathtubs or bath units, tanks and automotive parts. These enactments made people strongly aware that plastic waste recycling had become an important social problem.
and construction/furniture. The SMC Automotive Alliance, Southfield, MI, is an activity of the SPI’s Composites Institute, composed of 35 SMC plastics material molders and SMC raw material suppliers. One of their objectives is to prove that pyrolysis and grinding are viable methods of recycling automotive components made of SMC. The Center of Excellence for Composites Manufacturing Technology is operated for the Navy by the Great Lakes Composites Consortium, Kenosha, WI. One of the first projects undertaken was a prepreg (carbon/ epoxy) scrap reclamation project. The Big Three Vehicle Recycling Partnership of the US Council for Automotive Research has teamed with the Automotive Group of the American Plastics Council to create an infrastructure for plastics recycling of scrapped cars and trucks. Suppliers of Advanced Composite Materials Association (SACMA), Arlington, VA, is a trade association of composite material suppliers, parts fabricators and organizations that 40.10 ORGANIZATIONS ACTIVE IN provide ancillary support. Waste management, COMPOSITES RECYCLING including recycling composite waste, is one of The American Plastics Council, Washington, their areas of activity. ERCOM Composite Recycling GmbH, DC, is a joint initiative with the Society of the Plastics Industry Inc. (American Plastics Karlsruhe, Germany, is composed of SMC proCouncil, 1993). It was earlier known as the ducers and processors, resin suppliers and Council for Solid Waste Solutions. The Council glass fiber producers. It was founded to mission is to develop and implement a pro- develop and prove the recyclability of SMC in gram for the responsible use, recovery order to maintain SMC market share in light of (including recycling and energy recovery) and the strict German attitude towards waste manconservation of plastics. The Council is com- agement. Reinforced Plastic Waste Recycling and posed of 25 of the leading plastic resin producers, downstream customers and repre- Treatment Council, Chuo-Ku, Tokyo, Japan, a sentatives of the plastics processor committee of the Japanese Reinforced Plastics community. One of its four task forces is Society was organized to seek technical and Product Stewardship, devoted to advancing social solutions in disposal and reuse of large all options of integrated waste management. composite items such as ships, tanks and bath This task force, under the Durables Program, units. The Swedish Institute of Composites, Pitea, has identified four program groups to receive recycling (life-cycle management) attention. Sweden, is doing experimental work to show They are automotive, major appliance, com- that reground SMC can be incorporated into puter and business equipment, and building new SMC and BMC products.
904 Reuse and disposal ACKNOWLEDGEMENTS
The author wishes to acknowledge Mrs Peggy Renkel for her word processing expertise, helpful suggestions, diligence and patience in chapter preparation and Dr David Graham for his careful review of the text. REFERENCES American Plastics Council. 1993. Private Communication, T. Donnelly. What Industry Is Doing? Broudy, Phil. 1993. NIRS. Private Communication. Butler, Kurt. 1991. 46th Annual Conference, SPI Composites Institute, Session 18-B. Darrah, Basil. 1993. 48th Annual Conference, SPI Composites Institute, Session 15-B. DeMaio, Anthony. 1991. 46th Annual Conference, SPI Composites Institute, Session 18-C. Diesenhouse, Susan, 1994. Polyester Becomes Environmentally Correct. The New York Times, February 1994, Sec 3.5. Donnelly, Thomas 111. 1993. Status Report on National Legislation. SPE Recycling Conference,June 16. Chicago, Illinois. Graham, W. David, Jutte R.B. and Shipp, D.L. 1993. 48th Annual Conference, SPI Composites Institute, Session 15-H. Hanson, David. 1991. 46th Annual Conference, SPI Composites Institute, Session 18-E. Jutte, Ralph and W. D. Graham. 1991. 46th Annual Conference, SPI Composites Institute, Session 18-A. Kitamura, Tatsundo. 1993.48thAnnual Conference, SPI Composites Institute, Session 15-E. Lause, John. 1993. GLCC. Private Communication. McDermott, Joseph. 1992. Thermosets Recycled, What Now? Reinforced Plastics, 36(11). November. Miller, B. 1993. Recycling: What the Big 3 say. Plastics World, 51 (lo), 32-37
Nir, M., Miltz J. and Ram, A. 1993. Update on Plastics and the Environment. Plastics Engineering, 49,(3), 87-88. Norris, Donald. 1990. Proc. 1990 ASME/ESD Advanced Composites Conference. Olson, Barbara and H. DeKeyser. 1992. Recycling Cured Phenolic Material, Rogers Corporation, Molding Materials Div No. 5215-022-1.5E. Presented at SAE International Congress, Session Code M25B (February). Pettersson, Joakim and P. Nilsson. 1993. 48th Annual Conference, SPI Composites Institute, Session 15-E Protzman, Ferdinand. 1993. Germany’s Push to Expand the Scope of Recycling. The New York Times. July 4, sec 3%. Reitz, James. 1992. 47th Annual Conference, SPI Composites Institute, Session 4-B. Richter, Heinz and Jurgen Brandt. 1987. Recycling Carbon Fiber Scrap, Engineered Materials Handbook 7: Composites, pp. 153-156. Rusch, Ken. 1993. 48th Annual Conference, SPI Composites Institute, Session 15-G. Sahut, Jan 1994 Reinforced Plastic Making Some Big Fancy Parts. Plastic WorId, 52(10), 24 Schaefer, Peter and A. Plowgian. 1993.48thAnnual Conference, SPI Composites Institute, Session 15-D. SMC Automotive Alliance. 1991. 46th Annual Conference, SPI Composites Institute, Session 18-D. Tesoro, Guiliana et al. 1992. 47th Annual Conference, SPI Composites Institute, Session 4C. Tesoro, Guiliana and Y. Wu. 1993. Chemical Products from Cured Unsaturated Polyesters. Advances in Polymer Chemistry, 12(2), 185-196. Wigotsky, Victor. 1993. Plastics In Automotive. Plastics Engineering, 49(4), 36-38. Wood, J. 1991. A Disposal/Recycling Model for Composite Waste Materials. First International SAMPE Environmental Conference. May 1. pp. 388-390.
LAND TRANSPORTATION APPLICATIONS
41
Douglas L. Denton
41.1 INTRODUCTION
processes must be used to manufacture parts rapidly enough to meet the production rates of the assembly plants. As a result, injection and compression molding processes are used extensively for these applications. Composites typically used in transportation applications consist of low-cost grades of thermoplastic or thermoset polymers reinforced with E-glass fibers. Often these composites also contain mineral particulate fillers. Composites containing high-modulus fibers, such as carbon, and higher-performance resins such as epoxies, are used only where the higher cost can be justified to meet special product requirements. Metal matrix composites are used very sparingly in ground transportation. Over the past 50 years, the use of polymer composites generally has progressed from low-performance applications towards more demanding applications requiring excellent surface appearance, high mechanical properties, increased temperature stability or improved durability. With increasing demands to reduce vehicle weight for improved fuel economy, and to reduce investment costs for greater competitiveness,the use of composites in ground transportation is expected to increase for many decades to come.
The use of polymer composite materials in land transportation is steadily increasing because of the cost and performance advantages that composites offer over competing materials. Today, composites are extensively used in well-established passenger car, van and truck applications. New applications demanding higher structural performance levels are under development. Railroad cars, mass transit vehicles and a wide range of military ground transportation systems offer expanding opportunities for composite materials. Driving forces for the use of polymer composites in ground transportation applications include low manufacturing investment cost, cost reduction through parts consolidation, weight savings, good mechanical properties, 'Class A' surface quality, excellent durability characteristics, inherent dent and corrosion resistance, good noise and vibration dampening, styling flexibility and dimensional stability. Factors which tend to mitigate against their use are high materials costs, low modulus and the reluctance to use materials perceived to be 'new and unproven'. Only those materials and processes which provide required performance at the lowest cost achieve long-term commercial success in transportation applications. For high-volume 41.2 ECONOMICS AND MARKET GROWTH automobile and truck applications, high-speed The steady growth in the use of composites for land transportation is attributed primarily to Handbook of Composites, Edited by S.T. Peters. Published the development and acceptance of new applications. From the late-1960s to the mid-1990s in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
906 Land transportation applications be equal to or lower than stamped steel parts. A primary factor offsetting high materials cost is the lower investment needed to make parts from composites compared to steel. To produce a composite part generally only one mold and one press is needed whereas most steel parts require multiple stamping tools and presses to form the part. In addition, composites offer the opportunity to integrate several parts formed in steel into one part. This further reduces the number of tools and presses needed, and eliminates the welding operations required to join the stampings. The cost advantage achieved or cost penalty incurred by composites depends upon the total volume of an application. As the number of parts produced becomes very high, the reduced investment in tools and presses is offset by the higher materials cost for the composite. In the automobile industry the use of a composite part produced at a volume of several hundred thousand units per year may not be not cost effective. But production of the Year Fig. 41.1 Growth in the shipments of polymer com- same composite part at volumes of tens of posites for use in land transportation. (Source: SPI thousand units per year may be very competitive. The specific cost 'cross-over' point Composites Institute.) between composites and steel depends on a large number of factors and must be deterComposites have generally displaced metal mined for each specific application. alloys such as steel to gain new applications. In order for a material substitution to be suc41.3 HISTORICAL DEVELOPMENT OF cessful, composites must provide all of the APPLICATIONS functions required by the part at a competitive cost. Because of strong competitive forces in Polymer composites began to appear on cars the transportation industry original equip- shortly after World War I1 in small compoment manufacturers (OEMs) seldom pay nents under the hood and inside the higher cost for a new material to achieve desir- passenger compartment. The first major mileable functions beyond the part requirements. stone in exterior parts was the introduction of Therefore, composites must achieve cost effec- the Corvette in 1953 which sported body pantiveness for long-term use in transportation els made of glass fiber reinforced polyester applications. (Fig. 41.2). The body panels were produced On a price per unit weight basis, composite with open mold and preform molding material costs are generally several times processes up to 1972, when the production higher than the metals used in automobile and was converted to compression molded sheet truck applications. However, when the total molding composite (SMC). The invention of cost of component production and vehicle the SMC process and the development of 'low assembly is considered, composite parts can profile' polyester resins, which provided composite usage in USA transportation more than quadrupled (Fig. 41.1), and since the early 1980s it has expanded at a faster rate than the total composites market. Land transportation use of polymer composites represents the largest segment (over 30%) of the total market and reached 430 000 metric tons in 1994 according to the SPI Composites Institute.
Historical development of applications 907
Fig. 41.2 Glass fiber reinforced polyester body panels on the 1953 Corvette was the first major application of compositesin the automotive industry. Other composite applications introduced on the Corvette include bumper beams, leaf springs, radiator support, seat backs and rear floor pan. (Photo courtesy of General Motors Corporation.)
improved surface appearance, allowed the Fuel Economy (CAFE) standards led to intense proliferation of composites into higher vol- development of new composite applications and a significant upswing in composite usage ume body applications. SMC grille opening panels (GOP) were in the late 1970s (Fig. 41.1). Composites not introduced in the late 1960s and rapidly only expanded into additional cosmetic parts, spread throughout the industry because fiber but began to be seriously considered for use in glass GOPs saved weight and reduced cost structural components. Intense programs to through significant parts consolidation. With develop radiator supports, transmission supthe introduction of composite panels on auto- ports, leaf springs and wheels were initiated. mobile assembly lines, the dimensional Efforts to find additional under-the-hood and consistency of molded parts became well interior applications of composites were also known. Manufacturers also recognized that intensified. Improvements in composite surface quality the SMC panels could not be deformed to adjust for variations in the steel body struc- and productivity lead to the introduction of ture. Modification of assembly plant the 1984 Fiero, the second high-volume vehicle procedures to accommodate the unique char- with all exterior body panels made of composacteristics of composite parts continues to be ites. Using innovative body construction, exterior panels produced by the SMC and an issue with new part introduction. The need to reduce vehicle weight through RRIM (reinforced reaction injection molding) materials substitution and downsizing to meet processes were mechanically fastened onto a government mandated Corporate Average steel spaceframe. A similar design approach
908 Land transportation applications was used for the construction of General Motor’s all purpose vehicles (APV). Introduced in 1989, these minivans used approximately 120 kg (260 lb) of SMC per vehicle (Fig. 41.3). With annual production volumes ranging from 90 000 to 148 000 vehicles per year this represented the largest single application of composites in the automotive industry. The Dodge Viper was the first American vehicle to have all body panels produced by the resin transfer molding (RTM) process (Fig. 41.4). The low production volumes for the Viper (under 5000 per year) made this an ideal application for composites. The selection of an RTM process with low cost molds resulted in greater investment savings than could be achieved with SMC and RRIM. This milestone application was introduced in the early stages of an industry-wide effort to develop the RTM and SRIM (structural reaction injection molding) processes for the production of structural and Class A parts for the transportation market.
41.4 CURRENT APPLICATIONS
41.4.1 AUTOMOTIVE APPLICATIONS On average, cars produced in USA in 1994
contained about 50 kg (110 lb) of fiber reinforced composites. The broad usage of composites in the automobiles and trucks is illustrated by the list of applications in Table 41.1. Distinctions can be made between applications that are experimental, have been in limited production or are well-established in production. In many cases, breakthrough composite applications are introduced by manufacturers of low volume specialty vehicles before they appear in high volume car lines produced by larger companies such as Chrysler, Ford or General Motors. These larger companies often will evaluate a new application on a low-volume specialty vehicle or on a customized segment of a high-volume vehicle before committing to high-volume production with a composite part.
Fig. 41.3 GM’s all purpose vehicle (APV) is the automotive industry’s largest application of composites. Each vehicle uses approximately 120 kg (260 lb) of SMC in addition to other polymer composites. (Photo courtesy of General Motors Corporation.)
Current applications 909 7 -
-
I.
~ _ _.
-
-
~
~
-----zm
Fig. 41.4 The Dodge Viper is the first American car to have ail Class A composite body panels produced by the resin transfer molding (RTM)process. (Photocourtesy of Chrysler Corporation.)
Newly developed composites made from compression molded glass fiber mat thermoplastic Composites have proven to be very successful sheet (GMT) or liquid molded (RTM/SRIM) in a wide range of exterior body panels and thermoset resins may also compete with SMC are used in hundreds of vehicle applications. for horizontal body panels. Excellent surface finish, light weight, and a The broader use of composites in body panthermal coefficient of expansion near that of els is restricted by their limited coststeel have made these applications successful. effectiveness at very high production volumes. Customers appreciate the dent and corrosion Composite panels are well-suited for niche resistance of composite panels. vehicles and for providing economical opporSMC dominates composite applications in tunities to achieve styling differences within a horizontal panels such as hoods, roof and deck body line, but they are seldom employed in lids, and competes with RRIM for vertical pan- high-volume applications. Another factor that els such as fenders, doors and quarter panels. precludes the use of some polymer composites The use of unreinforced thermoplastic body in body panels is sensitivity to high temperapanels on vehicles such as G M s Saturn and ture. In most assembly operations all body Chrysler 's Intrepid, Concorde, and Vision panels are subjected to a heat soak at demonstrate the potential for the displacement 175-205°C (350400°F) for about 30 min folof composites in some body applications. lowing the electrodeposition of a corrosion
Automotive body applications
910 Land transportation applications coating (‘E-Coat’ or ’ELPO) to the steel body structure. Only a limited number of low-cost polymer composites can resist dimensional change or surface distortion during this heat treatment. While composites are used in the body structure for several low-volume sports cars, this opportunity remains essentially untapped in the auto industry. The Lotus Esprit has an all-composite body produced by the VARI process (vacuum assisted resin injection) that is mounted on a steel frame. Lotus uses a similarly produced composite floor pan and body panels in the Elan. Consulier manufactures sports cars with all-composite monocoque bodies produced by vacuum bag molding of epoxy prepregs with E-glass, Sglass, carbon and KevlaP fibers. The entire body structure weighs only 125 kg (275 lb) and takes all structural loads. Higher volume applications in body structure includes the Corvette which has used composites in the floor pan, rear underbody and radiator support. Since the early 1980s an evolving series of composite materials and processes have been used to produce front and rear bumper beams. SMC containingrandom chopped and/or continuous unidirectional fibers was initially used. Subsequently, stamped thermoplastic composites and SRIM bumper beams have been commercialized.Special design considerations and parts consolidationhave lead to the use of glass-reinforced polypropylene (Azdel@)in the front bumper beams of Chrysler minivans, which are produced in volumes of more than 500 000 units per year.
and extensive testing, a transverse rear leaf spring was introduced on the 1981 Corvette. This 3.6 kg (8 lb) filament wound E-glass/epoxy monoleaf spring replaced a ten-leaf steel spring weighing 19 kg (421b). A front transverse spring was added to the Corvette in 1984 and the first composite longitudinal spring appeared on the General Motors Astro van in 1985. The use of composite springs rapidly spread across GM car lines in the late 1980sand early 1990s. In less than a decade from its introduction, the Inland Division of GM (now Delphi Chassis Systems) was manufacturing more than one million Liteflex@composite springs per year (Fig. 41.5) to meet the demand.
Fig. 41.5 Composite leaf springs reduce vehicle weight, resist corrosion and outlast traditional steel leaf springs. (Photo courtesy of Delphi Chassis Systems.)
The exceptional weight savings and outstanding durability have been the keystones to the success of composite springs in cars and more recently, in heavy trucks. Composite springs Automotive chassis applications last at least five times longer than steel counWhile relatively few composite chassis compo- terparts in laboratory fatigue tests and in field nents have been commercialized, notable service. Added advantages are improved applications have provided evidence of the packaging due to smaller size, and improved performance level and durability offered by ride and handling characteristics. Their longcomposites. term field success has also clearly One of the most successful automotive stmc- demonstrated the survivability of polymer tural applications to date is the composite composite materials in the harsh under-vehileafspring. After decades of development work cle environment.
Current applications 911 Another suspension application demonstrating the durability and load carrying capability of composites is wheels. The commercial introduction of polymer composite wheels was on the 1989 Shelby CSX modified version of the Dodge Shadow. Developed by the Motor Wheel Corporation, the Fiberide@ wheel uses a combination of structural SMC and XMC. These materials provided a mixture of chopped random and oriented continuous E-glass in a vinyl ester resin. At reduced weight, the composite wheel outperformed both steel and aluminum wheels in fatigue. The development of polymer composites that retained acceptable lug nut torque after wheel heating caused by extreme braking conditions was a technical advance that made this application possible. A front suspension stabilizer link was introduced on the 1994 Ford Taurus and Mercury Sable. Made from a glass fiber reinforced copolymer polyacetal, this was the first use of a structural thermoplastic composites in such an application in North America. This durable part does not require painting and reportedly provides a 42% weight reduction and 33% cost reduction over the replaced steel part.
by the 'lost-core' process in which a low temperature melting alloy such as tin-bismuth is used as a form to mold the hollow sections. The core is then melted away from the cured part in a subsequent operation. An alternative is to 'weld' two thermoplastic pieces to form the manifold. To reduce cost, some fuel intakes are being integrated with the air intake manifold. Injection molded phenolic compounds are used in many engine and transmission applications because of dimensional stability and creep resistance at higher temperatures. Current applications include pulleys, torque converter reactors, thrust washers, water outlets, valve covers, radiator end caps, motor commutators and fuel rails (Fig. 41.6). While the performance advantages of composite drive shafts and propeller shafts (lower weight, better NVH, increased durability) has been established, cost remains a barrier to their wide spread use. A composite drive shaft has an economic advantage, however, when it replaces two-piece metal shafts. The most successful designs have used glass and/or carbon fiber composite overwrapped on an aluminum tube using pultrusion or filament winding processes.
Automotive powertrain applications The application of composites in the powertrain is extensive and is growing. In the past, only metals were considered for use in the demanding environment of the engine and transmission. Increasingly, composites are being selected to reduce weight and cost, and to improve NVH (noise, vibration and harshness), engine efficiency, packaging and corrosion resistance. Air intake manifolds are rapidly being converted from metal to injection molded thermoplastics reinforced with short glass fibers. Introduced on European cars in the 1970s, composite manifolds are predicted to be used on most vehicles by the turn of the century. Most intake manifolds are produced
Fig. 41.6 Powertrain applications of phenolic composites include transmission torque converter reactors and thrust washers, poly-V pulleys, and water outlets. (Photo courtesy of Rogers Corporation.)
912 Land transportation applications Experimental engines made almost exclusively of polymer composite materials, including block, head, pistons and connecting rods, have been tested, but commercialization in the near future. Many is matrix (MMC) parts have been prototyped for engine applicationsf few have gone into production (Table 41.1). Aluminum reinforced with silicon carbide, alumina or carbon fibers offers the potential for good mechanical and thermal properties with large weight savings. However, factors such as high cost, low ductility and machining difficulties have retarded the commercialization MMCs in transportstion components. Automotive interior applications
41.4.2 TRUCK APPLICATIONS
Composites are widely used in the mediumduty and heavy-duty truck industry to decrease weight, reduce manufacturing and maintenance costs, and extend the longevity of the vehicles. Lighter trucks allow increased payload capability and improved fuel economy. Since trucks are generally not produced in high volumes, the lower tooling invesment favors composite parts. The durability is important since are used for many years. Composites are commercially used for essentially all skin surfaces on truck cabs, air deflectors, hoods, fenders, roofs, side closure panels, sleeper box and doors. Mack Truck introduced the first structural SMC door in 1978 and two years later GM Truck and Coach made the first use of continuous fiber SMC in the door of the Astro and General truck cabs. In 1983 Mack Truck became the first manufacturer to introduce a cab with composites used in all exterior surfaces. These panels covered a frame made of steel, aluminum, and structural composite beams. Several all-composite truck cabs using monocoque structures have been prototyped and shown to meet all performance requirements. Weight savings of 20% are achieved in truck trailers when floor, wall and ceiling panels are composite. In addition, refrigerated trailers made from composites are 25-30y0 more thermally efficient than metal trailers. In a recent application, Trail King has introduced a flat bed trailer with a fiber reinforced epoxy deck to achieve lower weight, corrosion resistance and lower cost. One of the fastest growing application in heavy truck is composite leaf springs for large weight savings and greatly extended spring life.
manufacturers make extensive use of polymeric materials on the interiors of cars and trucks/ most do not require fiber reinto meet performancerequirements. Some notable exceptions are seats, load floors, knee and instrument Pane' seat were first introduced on the 1975 Corvette. In addition to stamped polypropylene, thermosetting Polyurethanes and Polyester are used inboth seat back and seat pans. An injection seat with long fiber reinforcement Premiered On the 1993 Dodge sheet is currently preferred for load floors in station wagons and extended pick-up truck cabs. Stamped polypropylene has been used in numerous knee bolsters. This structural part at the bottom of the instrument panel helps to control occupant movement during a frontal crash. Composites and unreinforced thermoplastics are also being used in instrument panel supports to provide lower cost through parts integration and reduced weight. The use of composites in the front of vehicle interiors is 41.4.3 RAILROAD APPLICATIONS expected to expand. Durability and light weight make composites attractive for railroad applications, yet few
Current applications 913 Table 41.1 Composite applications in automobiles and trucks' Experimental
Limited production
Well established
Body
Front rails A, B and C pillars Roof frame Rockers Rear underbody Pickup truck box Truck cab structure Truck trailer bed
Radiator support Floor pan Cowl panel Wheel housing Windshield surround Door surround Center tunnel Tanker truck trailers EV battery tray
Grille opening panel Grille opening reinforcement Rear end panel Hood Deck lid Lift gate Rear hatch Roof panel Hardtop cover Fenders Quarter panels Doors outers Door inners Spoiler Headlamp cover Fuel filler doors Front bumper beams Rear bumper beams Truck air defector Truck trailer walls Truck trailer roof
Chassis
Front cross member Transmission support Brake rotors'
Wheels Stabilizer bar links
Leaf spring Disk brake pistons
Powertrain
Throttle body Oil pump Transmission valve body Hydraulic clutch actuator Engine - Pistons, block, head, piston pins3 connecting rods3 rocker arms2 bed plate2 Flywheel4 Drive shaft2
Drive shaft Propeller shafts Water pump housing Water pump impeller Fuel pump components Fuel rail Fuel tank supports Camshaft sprockets Oil pan Differential cover Valve lifter guides CNG storage cylinders5 Cylinder liners2 Pistons2
Air intake manifold Battery tray Fan shroud Radiator end caps Transmission torque converter reactor/stator Transmission thrust washer Motor commutators Diesel electronic unit injector Valve cover Timing chain covers Poly-V pulleys Idler pulley Distributor cover
Interior
Car jack Steering wheel
Knee bolster Instrument panel support Window frame/trim
Seat back Seat support Load floor Plenum Glove box
Polymer composite unless noted otherwise Metal matrix composites Metal matrix or polymer matrix composites
Electric powered vehicles Natural gas vehicles
914 Land transportation applications have been commercialized. Pultruded glass fiber reinforced thermoplastic composite panels are used to construct a 'Secured Modular Automotive Rail Transport' (SMART) for Union Pacific Railroad. Each three-tier structure serves as a 'car rack' to protect eighteen automobiles from damage and theft during transport to the dealer. The use of composites also significantly reduces the maintenance cost of the car rack. Composite hopper car covers protect grain and other dry materials that need protection from moisture. The covers, which range from 9-15 m (3040 ft) in length, are fabricated by a hand lay-up process. Two prototype glass/polyester filament wound railroad cars termed the 'Glasshopper' have been in service since 1981 without failure. Produced in a joint venture by ACF Industries, Cargill and Southern Pacific, the corrosion resistant, lightweight composite car can carry a greater payload, but is significantly more expensive than a steel hopper car. 41.4.4 MASS TRANSIT APPLICATIONS
Buses and passenger rail systems offer many more opportunities for the application of composites than are presently in service. Some cosmetic and semi-structural applications have been successfully implemented, but few examples of structural components are found. The durability of polymer composites has lead to extensive use in seats for buses, subways, people movers and trains. Composite sandwich panels with glass fiber-phenolic skins over aramid or aluminum honeycomb are used in walls, ceilings and floors of many European mass transit cars. These rigid panels are very lightweight and the use of phenolic resins allows attainment of fire/emission standards. The end caps of transit cars are often molded composites. Pultruded exterior panels are being substituted for aluminum on buses to reduce weight and decrease assembly cost. The development of high speed rail systems, which attain speeds up to 480km/h
(300 miles per hour), may offer an opportunity for composites because of the need to minimize weight. Weight reduction is especially important for magnetic levitation (MAGLEV) systems where the vehicle is suspended above the guideway to provide friction free movement. 41.4.5 MILITARY APPLICATIONS
Limited applications of composites are found in combat and non-combat ground vehicles in the US military, but if current development programs are successful, much greater use of composites should result. Modern warfare requires rapidly deployable, survivable vehicle systems. Thus, weight reduction in all vehicles, including armored vehicles, is desirable. In addition to weight savings, composites can potentially offer increased durability, improved signature management! better personnel protection and lower production cost. The goal of an ongoing Army project, the Composite Armored Vehicle Advanced Technology Demonstrator (CAV ATD), is to establish the feasibility of using polymer composites in the primary structure of a 20 tonne (22 ton) combat vehicle to achieve a 33% weight saving over a traditional metal vehicle. Composites are also being considered for combat vehicle armor - used either alone or in conjunction with ceramics and/or metals - to provide significant weight savings over current materials. The High Mobility Multi-purpose Wheeled Vehicle (HMMWV), which is currently in production, utilizes an integrated hood/fender assembly molded from SMC. In upgraded versions, composite armor is attached to the HMMWV. Phenolic spa11 liners containing Kevlar or S-2 Glass@fibers are used in the M113A3 Armored Personnel Carrier and Bradley Fighting Vehicle to provide troop protection. Longer service life and chemical protection were the motivation for the incorporation of composite seats and side racks in 2.2 tonne
Conclusions 915 (2.5 ton) and 4.4 tonne (5 ton) trucks. Other and technical expertise acquired in defense potential composite applications such as M1 programs to the development of composite Abrams Main Battle Tank components (dri- technology needed to improve the competiver’s seat, air intake plenum, stowage box, tiveness of American industries with a and power pack container), tactical vehicle particular emphasis on the automotive indusbody, fuel tanker and HMMWV drive shaft try. Numerous federal programs are directing are under consideration for production after dollars and technical expertise resident in the the turn of the century. National Laboratories into ground transportation programs. A notable example is the ’Partnership for a New Generation of Vehicles’ 41.5 FUTURE DIRECTIONS program, which teams the government with The growth of the transportation market is the automotive industry to produce the techexpected to continue, and potentially acceler- nology to make safe, comfortable, and ate, into the 21st century. As composites affordable cars that achieve up to 30 km/l (82 become better understood by designers, and miles per gallon) of gasoline with low emisas the reliability and advantages of these mate- sions. Composite materials are expected to rials are more clearly demonstrated, they will play a key role in meeting this challenging be used increasingly in more demanding goal. Successful implementation of the techstructural applications. Body structure, chassis nologies developed in these cooperative and powertrain offer tremendous opportuni- programs could have an enormous impact on ties for the utilization of composite materials. the usage of composites in future vehicles. Improved manufacturing capability to rapidly produce composite parts will boost their eco41.6 CONCLUSIONS nomic viability in high volume applications. The development of a commercial infrastruc- Ground transportation is the largest and one ture and market for the large scale recycling of the fastest growing segments of the polymer and reuse of polymeric materials will also composites market. While transportation use increase the growth opportunity for compos- of composites is expected to expand, the rate ites. of growth depends on a number of factors. The potential economic advantages and Improvements in technology are needed to weight reductions afforded by the use of com- increase high volume production capability posites in integrated body structure and and increase the cost competitiveness of comchassis components has driven development posites relative to other materials. The programs since the 1980s. In 1988 Chrysler, performance and durability of composites in a Ford and General Motors formed the wide range of structural applications must be Automotive Composites Consortium (ACC) to demonstrated. More designers and engineers conduct joint R&D on structural polymer com- need to become familiar with the unique charposites with a focus on these structural acteristics of composites and learn to develop applications. Operating under the United States designs that use the full potential of these Council on Automotive Research (USCAR),the materials. An economically viable infrastrucACC works with supplier companies and uni- ture for dealing with post-consumer waste versities to develop the processing, materials, must be established. With these advances, design and joining technology needed to composites are expected to play an important achieve production worthiness and cost effec- role as industry meets the increasing worldtiveness of composite structures. Currently, the wide demand for safe, clean, energy efficient US Federal Government is providing resources land transportation.
MARINE APPLICATIONS
42
Wayne C. Tucker and Thomas Juska
resins, but their superior resistance to hydrolUse of composites in marine applications is ysis and blistering aacquemet and LaGrange, widespread. The two major advantages of fiber 1988) makes isophthalic polyester a better reinforced dastics over metals are resistance to choice for applications requiring long term the marine environment, particularly the elim- exposure to water. Use of vinyl esters is becoming more comination of galvanic corrosion and the ease of mon. Although they are more expensive than tailoring structures, which are fabricated by polyesters, vinyl ester laminates generally molding processes. In addition, composites In addition, vinyl esters have better properties. have high strength-to-weightratios. have excellent resistance to matrix hydrolysis. This chapter is an overview of the materials Also available are mixtures of vinyl ester and and fabrication processes used in marine applipolyester, which offer some of the benefits of cations of composites. More comprehensive vinyl ester but at intermediate cost. studies of the use of composites in marine conEpoxy laminating resins, rarely used in struction have recently been published (Smith, marine applications due to cost and the need 1990;Greene, 1990; Davies and Lemoine, 1992). for elevated temperature cure, are used by some boat builders in high performance one42.2 MATERIALS off racing craft and some production boats. Among the advantages of epoxy are extended 42.2.1 RESINS out-time compared to room temperature curing resins and low volatile organic content. Resins used in marine applications generally The material form is wet epoxy prepreg cure at room temperature, both for the low (sometimes referred to as a wet-preg to distinfabrication costs and because elevated temperguish it from conventional prepreg) made by ature performance is not required. the boat builder (Gougeon, 1992). Fabrication General purpose polyester is the most comis normally by vacuum bag, room temperature mon laminating resin. These materials, based cure followed by oven post-cure at about 50°C on orthophthalic acid (phthalic anhydride), (122°F). are the least expensive, but long term immersion without a barrier coat will probably result in blistering (Burrell, Herzog and McCabe, 42.2.2 REINFORCEMENTS 1987). Polyesters based on isophthalic acid are slightly more expensive than general purpose E-glass fabric is the primary reinforcement in marine construction, of which there are numerous forms. There are two basic styles: woven Handbook of Composites. Edited by S.T. Peters. Published and knitted. Woven fabrics are further subdivided into woven roving and woven yarn. in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7 42.1 INTRODUCTION
Materials 917 Woven roving are the most common fabrics due to their low cost. They are available in a variety of weights and weave patterns, but a 800 g/m2 (24 oz/yd2) plain weave is probably the most frequently used. In small boat construction, the fabric is usually mat-backed, in which a chopped strand mat has been stitched or powder bound to one side of the fabric. Woven yarn, also called textile fabric or cloth, is more expensive and therefore used infrequently. These materials are finer in texture than woven rovings and are used as surfacing plys, particularly in tooling. Knitted glass fabrics are becoming more common. The rovings are stitched together, which keeps the strands relatively parallel (no crimp as in weaves) and stabilizes the fabric against distortion during handling. The orientations of the separate layers and the number of separate layers, which compose knitted fabrics can be varied, although a 2-ply 0/90 is most common. Knitted fabrics are currently available up to 2500 g/m2 (72 oz/yd2). Chopped strand mat is most commonly used where a resin-rich layer is desired, such as between the core and fabric in sandwich structures. Usually the first ply next to the gel coat is a layer of mat, which reduces print-through of the reinforcement as well as providing additional protection from water permeation. S-2 glass is about six times the cost of Eglass, so it is rarely used in marine structural applications. S-2 made in the G filament is sized for epoxy and when made with the K filament is sized for polyesters and vinyl esters. Carbon fiber is also rarely used because of cost. America’s Cup racing sailboats use carbon throughout the boat and some racing powerboats use carbon as well. Carbon has an additional problem in that it can corrode metals in contact with it in the presence of water (Tucker, Brown and Russell, 1990; Aylor and Murray, 1992). Polymeric fibers, namely Kevlar and Spectra, are sometimes used in marine applications where weight is critical (Sloan and
Nguyen, 1992).Their use is limited because of cost and also because they cannot bear much load in compression. 42.2.3 CORES Linear PVC foam is used extensively in boat hulls, at a density of 80 kg/m3 (5 lb/ft3). The resilience of the material allows hull deformation during impact with no skin/core debonding, skin damage or core shear failures. Crosslinked PVC foam is used in decks, bulkheads and cabins because it has better structural properties and will not creep at the higher topside service temperatures. It is available in a wide range of densities, but is usually used at 80 and 100 kg/m3 (288 and 3601b/yd3). Both types of PVC foam are closed cell, absorb very little water and are moderately priced. End grain balsa, also widely used as the core in boat sandwich structures, is inexpensive and has excellent structural properties. Under some circumstances, however, balsa can rot. If skin/core debonding occurs and water is allowed to penetrate this interface, the wood will eventually degrade. It has been shown, however, that balsa will provide a long marine service life with proper maintenance (Baltek Corporation, 1986). Composite core is also available. The material is made in two forms, either sinusoidal nested-wave, or in a closed cell configuration where planar webs alternate with sinusoidal webs (Plunkett et aI., 1992).This type of core is essentially a collection of sine wave stiffeners fused to skins with a laminating resin. It appears to have great potential for marine and civil engineering structures and has in fact been used to make about 200 prototypes, including deckhouses, storage tanks, buildings and parabolic solar energy concentrators. 42.2.4 SKIN/CORE BONDING
A good bond between skin and core is critical to the performance of sandwich structures. In
918 Marine applications boat building it is general practice to use vacuum bag pressure to achieve the necessary quality. There are two basic methods: one is to use a paste adhesive between skin and core and the other is to use the laminating resin as the adhesive. Paste adhesives for a polyester or vinyl ester laminate are filled polyesters. The adhesive is catalyzed and troweled onto the surface of precured laminate (normally the outer skin) and the foam core is sprayed with catalyzed resin (Johannsen, 1990). The core is then bedded into the paste, the vacuum bag is positioned and the assembly is allowed to cure at room temperature. The core either has holes to allow air to escape, or contour core is used. In both cases, observation of the paste seeping through the spaces is evidence for a successful bonding operation. The inner skin is laminated directly onto the surface of the core, with the first ply normally a chopped strand mat. Some sandwich structures are made by bonding the core with a resin rich layer. In this method, a wet-out ply of chopped strand mat is applied to the outer skin and used as a bedding for the core. In a recently discovered variation of this approach, which appears to offer some advantages, a ply of polyester mat is used to bed the core (Lewit, 1990).
Fig. 42.1 Boat hull blisters. (Courtesy of Dr. Tom Rockett, University of Rhode Island.)
3. More water is drawn into the laminate through osmosis, creating a localized pressure. The solution in the osmotic center hydrolyzes the resin, resulting in an increase in water-soluble materials, which then draw more water into the laminate and increase the pressure.
There are materials solutions to the blistering problem of orthophthalic polyester laminates. One is to coat the hull below the waterline with an epoxy paint formulated to reduce water penetration (Interlux, 1993). Another is 42.3 ENVIRONMENTAL EFFECTS to use a permeation barrier between gelcoat and laminate, as mentioned previously. A 42.3.1 BLISTERING more expensive solution to blistering is to change to a vinyl ester laminating resin. Use of Blistering can occur in gel coated marine lamisophthalic polyesters is another alternative, inates (Fig. 42.1). The causes are complex and which offers a compromise between cost and blisters may occur due to the selection of materesistance to hydrolysis. rials or due to the process used in fabrication Blistering of carbon fiber reinforced plastics or both. The formation of blisters basically (CFRP) in seawater can occur for another reainvolves three steps (Marino, Rockett and son. When galvanically coupled to a metal Rose, 1985): fitting, the CFRP becomes cathodic. The result1. Water permeates into the laminate. ing electro-chemical reaction produces 2. Water-soluble components in the resin are hydroxyl ions near the carbon fibers. To baldissolved in this water and are concen- ance the charge, sodium ions are drawn in. trated in cracks or voids in the laminate, The eventual result is blisters with high alkacreating an osmotic center. linity (Tucker, 1991; Miriyala et al., 1993).
Fabrication 919 42.3.2 MICROBIAL DEGRADATION
production rates. These machines perform All engineering materials become colonized three important functions: they wet out the by microorganisms within hours after expo- fabric, meter the resin and place the wet out sure in natural waters (Little and Wagner, fabric via an overhead crane or gantry 1986). Microorganisms grow and produce a (Venus-Gusmer, 1989). Laminators must still viscoelastic layer or biofilm. The environment roll out the fabric after it is placed to remove at the interface is radically different from the air. Resin contents are reduced to about 50 bulk medium in terms of pH, dissolved oxy- wt %. Portable impregnators are also occasionally gen and organic and inorganic species (Little et used. Some high performance racing sailboats ai., 1991).In addition to biofouling, the materare laminated with wet epoxy prepreg made ial properties may be adversely affected. by the fabricator. Epoxy and nylon coating on steel can be breached by mixed cultures of marine bacteria (Jones, Walsh and Mansfield, 1991). Sulfate 42.4.3 RESIN TRANSFER MOLDING reducing bacteria (SRB) degrade marine caulks and polymeric coatings (Jones, Walsh Vacuum assisted resin transfer molding and Mansfield 1991; Jones et ai., 1992a,b). (VARTM)has been shown to have great potenOrganic surfactants on S-2 glass fibers have tial for the fabrication of ship hulls and reportedly been attacked in the presence of secondary structures. The Naval Surface SRB as well (Tucker et al., 1993). Warfare Center, Carderock Division, has had Microorganisms can be a threat to material several prototype structures made with a VARTM method known as SCRIMP (Seemann, performance in the marine environment. 1989), including deckhouse modules, masts and a Swimmer Delivery Vehicle (NSWC Carderock Division, 1993). The advantages of 42.4 FABRICATION VARTM over hand lay-up are that very low resin contents are achieved (30-35 wt YO), the 42.4.1 HAND LAY-UP process is less labor intensive and styrene Hand lay-up is the most common fabrication emissions are almost eliminated. Laminate method in marine construction. Catalyzed mechanical properties of panels made by resin is first applied to the gel coated mold VARTM show significant strength and stiffsurface, usually by spraying. A layer of dry ness improvements over panels made by hand mat or fabric is positioned onto the resin and lay-up (Juska,Mayes and Seemann, 1993). by pressure of hand-held rollers, resin gradually soaks through the fabric and most of the trapped air is forced out. This process is 42.4.4 VACUUM BAG PROCESSING repeated until the desired number of plys is Vacuum bag processing is used extensively in attained. The full thickness of the structure is the boat building industry to facilitate bonding rarely built up all at once, but is usually done of the core onto pre-cured skins, as described in several lay-up/cure stages. Resin contents earlier. Vacuum bags are rarely used to compact are normally about 55 wt % and void contents polyester or vinyl ester wet lay-ups because about 1%. there is little time to place the bag before gelation. Vacuum bag consolidation is used, however, to laminate wet epoxy resins and 42.4.2 IMPREGNATORS oven-cured prepregs, but the practice is limited Impregnators are sometimes used in the fabri- to the production of high performance boats cation of large composite structures to increase such as the America’s Cup racing sailboat.
920 Marine applications and can be used as a guide to the properties of typical marine laminates (made by hand layup) with those attainable by alternative fabrication methods. The properties of cored panels are becoming increasingly important to the design of marine structures, particularly boats. Recent studies have indicated that cored panel prop42.4.5 FILAMENT WINDING erties as determined with standard beam Filament winding is very often used to make bending tests should be supplemented by structures and machinery for marine applica- measuring the response of plate specimens tions. It is inexpensive, largely automated and subjected to uniform pressure loading (Huss, low resin contents are achieved. It is most 1990; Reichard, 1992; Gougeon and Bertelsen, appropriate for parts with a circular cross sec- 1993). tion, such as shafting, piping and pressure bottles.
Autoclaves are rarely used in marine construction, mainly because the process is too expensive for marine structures. There are a few exceptions, however. Submarine sonar bow domes are autoclave cured, as are carbon/epoxy sailboat masts.
42.6 STRUCTURES
42.5 MECHANICAL PROPERTIES
42.6.1 MINE COUNTER MEASURE VESSELS
Materials with a wide variety of fibers, resins, resin weight contents and void contents are used in marine construction. Extensive testing at NSWC, Carderock Division, has led to a database of marine construction materials made by hand lay-up, VARTM and vacuum bag consolidated prepregs ouska, Ma yes and Seemann, 1993; Juska 1993). The values in Table 42.1 were abstracted from this database
Mine counter measure vessels are among the largest GRP boats fabricated. These military vessels are made of non-magnetic material (GRP or wood) to reduce the threat from magnetic mines. Several foreign Navies have built composite minehunters, including the Soviet, British, Swedish, Italian, French and Australian, usually 4548 m (150-160 ft) long (Greene, 1990).
Table 42.1 Properties of marine laminates Fabric resin Fabrication
Woven roving vinyl ester
Woven Style 7781 roving epoxy vinyl ester
Hand lay-up
VARTM
Prepreg
Resin content (wt Yo)
50-60
30-35
35-40
Void content (%)
0.5-1.5
0-0.5
0.54
Compression strength, MPa (psi x lo3)
172-207 (25-30)
379414 (55-60)
379414 (55-60)
Tensile strength, MPa (psi x lo3)
172-207 (25-30)
379414 (55-60)
379414 (55-60)
Tensile modulus, GPa (psi x lo6)
11.7-13.8 (1.7-2.0)
21.4-22.8 (3.1-3.3)
21.4-22.8 (3.1-3.3)
Flexural strength, MPa (psi x lo3)
276-345 (40-50)
483-552 (70-80)
483-552 (70-80)
Structures 921 The US Navy, whose Mine Counter Measure (MCM) ships have a wooden hull construction, is currently building the MHC51 class coastal minehunter with GRP (Hepburn, Magliulo and Wright 1991). The Navy selected to modify the Italian Lerici class design for the MHC-51, which has an unstiffened single skin hull. The resin is an isophthalic polyester formulated to have a long gel time (about four hours). The reinforcement for the hull, decks and transverse bulkheads is DF1400 (1400 g/m2; 40 oz/yd2) composed of spun roving in the fill direction. The superstructure is constructed of Rovimat 1200, an 800 g/m2 (24 oz/yd2) woven roving stitched to a 400 g/mZ (12 oz/yd2) mat. The reinforcement is wet-out and placed with an impregnator.
42.6.3 LARGE POWER YACHTS
Large custom private power yachts, typically 21-39m (70-130 ft) long but up to 48m (160ft), are made with GRP by several boat builders. Most of the builders make hulls cored with linear PVC, with decks, bulkheads and superstructures cored with crosslinked PVC. The cores are bedded in paste adhesive and vacuum bagged to the outer skin. Blister protection is usually provided by several layers of epoxy paint applied to the gel coat after the boat is demolded. Westport Shipyard developed the concept of using the same (female)mold to make several non-identical boats. They have pulled 35 hulls from a single mold, which they alter in length and width to accommodate the design. Westport uses an overhead crane impregnator, which they believe saves 50% in labor over hand lay-up. They laminate with a general 42.6.2 SMALLBOATS purpose polyester resin and a 800g/m2 Small boats and craft are usually constructed (24 oz/yd2) woven roving (not mat backed). from GRP and probably dominate the marine Other builders have also adopted the varicomposites market. There are numerous styles able mold approach. The Christensen Motor of power and sail boats. In spite of the diver- Yacht Company makes boats by hand lay-up sity in hull shape and size there are fairly up to 43m (140 ft) long. They prefer knitted standard fabrication techniques and materials reinforcement (2408) for its handling characused in boat building. A comprehensive teristics and more uniform resin content and review of the methods and materials used in use vinyl ester in the outer skin. North Coast Navy craft fabrication and repair has recently Boats has a mold capable of producing boats been published (Russell, 1992). up to 40 m (131 ft) long. They make boats by Female molds are more common than male hand lay-up with a relatively heavy knitted plugs in production boats due to the impor- reinforcement (3205) and a vinyl ester/polytance of unvarying hull shape and ester resin mixture. The Nordlund Boat appearance. After the mold surface is coated Company laminates with a 600 g/m2 with release agent, it is sprayed with several (18 oz/yd2)knitted reinforcement for the outer layers of gel coat (usually an isophthalic acid, skin, for reduced print through and woven neopentyl glycol polyester) until a roving in the inner skins. 0.5-0.76 mm (20-30 mil) wet thickness is Some builders of large yachts do not use a achieved. The gel coat is allowed to harden variable mold. Admiral Marine specializes in prior to lamination. The first ply of the lami- one-off female tooling. They used this concept nate is usually a chopped strand mat to to fabricate a 48 m (160 ft) private motor yacht reduce print-through of the subsequent layers (by hand lay-up). An E-glass/Kevlar hybrid of woven or knitted reinforcement. Boats are fabric was used in the decks and cabin. usually fabricated by hand lay-up, as Jones-Goodell Yachts also make a unique described earlier. female tool for each boat, producing boats UP to 33 m (110 ft) long.
922 Marine applications Delta Marine produces displacement, semi- typical of marine construction, however; since displacement and planing hulls using a the designs minimize weight and maximize variety of techniques. The displacement hulls stiffness of structural components, the yachts are built using a stiffened single skin, where use aerospace materials. For example, the the semi-displacementhulls in the larger sizes boats built for America3Foundation's successutilize a balsa cored sandwich construction in ful defense of the 1992 Americas Cup had the sides and a stiffened skin bottom. The pri- carbon/epoxy tape in the hull (cored with mary materials used are a mat backed knitted aluminum honeycomb), frames, keel, mast, fabric (3205) and a combination of vinyl ester pole and boom, carbon winch drums and and general purpose polyester laminating shafting and pultruded Kevlar standing rigresin. In addition to private motor yachts, ging (Kramers, 1993). The design philosophy Delta builds commercial fishing vessels and was to keep the weight as low as the IACC rules allowed, while maximizing stiffness and small charter boats. being just strong enough to avoid failure. The processing conditions for the hull are limited 42.6.4 AMERICA'S CUP RACING SAILBOATS to 1atmosphere pressure (i.e. vacuum bag) at The America's Cup class yachts use compos- 95°C (203°F))while those for the mast are 3 x ites more extensively than any other marine lo5 N/m2 (3 atm) and 135°C (275°F). structure. The materials in the boats are not MAIN THRUSTERS
RECOVERY STROBE LIGHT
\
ELEVATORS
\RF BEACON TITANIUM HEMISPHERE
RECOVERY
RETRACTABLE LIFTING EMERGENCY
GRAPHITE COMPOSITE PRESSUREHULL
DESCE E !E Z I::
/
\\
FORWARDLOOKING OAS SONAR CCD CAMERA
\
FORWARD VERTICAL THRUSTER
\
SIDE-LOOKING SONAR
' I
DOPPLER TRANSDUCERS
NAVIGATION TRANSPONDER
AFT VERTICAL THRUSTER
Fig. 42.2 Advanced Unmanned Search System. (Courtesy of Naval Command and Ocean Surveillance Center, RDT & E Division.)
Structures 923 42.6.5 PRESSURE HULLS
42.6.7 FAIRINGS
The potential for composites in pressure hulls has been demonstrated (Garvey, 1990; Stachiw and Frame, 1988).Actual use of composites in these structures has been limited, however. One exception is the Advanced Unmanned Search System (AUSS), developed and operated by the Naval Command, Control and Ocean Surveillance Center, RDT&E Division (NRad, 1992). It consists of a vehicle, a control van, a maintenance van, a launch and recovery ramp and tow fish for acoustic communications. As shown in Fig. 42.2, the vehicle has a graphite/epoxy cylindrical pressure hull with titanium hemispherical endclosures. The design depth is 6095m (20000 ft) and it has had 134 successful operations, including dives to 3557 m (12 000 ft). The AUSS locates and inspects objects on the ocean bottom with sonar and cameras and transmits the information to the surface acoustically. It is an untethered vehicle, receiving commands through the acoustic link.
The submarine fairwater, or sail, (Fig. 42.3) is an excellent example of a large complex composite marine structure. A typical fairwater is 12m (40 ft) long, 73m (24 ft) tall and composed of over 1000 component sections. These massive structures were retrofitted on numerous classes of submarines replacing steel fairwaters. The composite fairwaters provide the Navy with lighter weight and non-corrosive properties that steel could not match. The construction of submarine fairwaters consists of E-glass/polyester panels and secondary stiffeners bolted to the metal superstructure. Mast fairings on Navy submarines are fitted
42.6.6 SONAR DOMES
There are numerous submarine and surface ship sonar domes. The largest sonar domes in service for the Navy are submarine bow domes. All bow domes made to date were fabricated at the Terminal Island facility of Hitco’s Defense Products Division. The structures are fabricated with glass/epoxy prepreg and are autoclave cured in one stage, which is remarkable because of the size, weight and thickness of the parts. The material characteristics necessary to allow the one-stage cure are stringent and only four materials have been qualified for use in sonar bow domes (Hitco, 1986). The mechanical properties specified are also restrictive, particularly the value of Mode I fracture toughness, where an incipient value of 2627 J/m2 (15 in lb/in) is specified.
!
:
I
Fig. 42.3 Submarine fairwater. (Courtesy of Lunn Industries, Inc. Wyandanch, NY.)
924 Marine applications
on periscope masts, antenna masts and snorkel masts. The masts are constructed from epoxy/fiberglass and some types are filament wound. Figure 42.4 illustrates a propulsion shaft fairwater for the US Navy’s CG-47 class cruisers. These hybrid-composite components were filament wound with continuous carbon and glass fibers on a six-axis filament winder. They provide a 50% reduction in weight over their metallic counterpart while exhibiting superior stiffness and strength. These cantilevered structures are subjected to large hydrodynamic loads. The composite system resists corrosion and biofouling due to a co-cured external jacket which also improves acoustic characteristics.
42.6.8 CONTROL SURFACES
The NR-1 nuclear powered research submarine was targeted by NAVSEA and the David Taylor Research Center as a test platform for the first all-composite diving planes. The diving planes are hybrid composite construction with a syntactic foam core. 42.6.9 OILPLATFORMS
The offshore oil industry is making use of composite materials on their platform structures. Examples of applications are: composite drain lines, composite cable trays, composite well head access platforms (Fig. 42.5) and composite firewater systems (Fig. 42.6). The structures
Fig. 42.4 Propulsion shaft fairwater. (Courtesy of B.F. Goodrich Aerospace, Engineered Polymer Products Business Group, Jacksonville,FL.)
Machinery 925 42.6.11 DECKING Aside from the application mentioned in the section on oil platforms (Fig. 42.5), composite decking is used where lightweight non-corroding decking is needed. Examples are ship superstructures, work platforms and catwalks. 42.6.12 SHIPBOARDARMOR Composite deckhouse armor is used by the Navy in the LHD class. LHD-1, the USS Wasp, has panels composed of Kevlar reinforced Fig. 42.5 Composite wellhead access platform. vinyl ester. For LHD-2, LHD-3 and LHD-4, the (Courtesy of EXXON Production Research material was changed to S-2 fabric-reinforced Company.) phenolic, each vessel of which carries 64 000 kg (140 000 lb) of armor plate (Owens Corning Fiberglas, 1991). The change was made to reduce cost without sacrificing ballistics protection. The fiber/matrix adhesion in the S-2 material was deliberately engineered to I' be poor, which results in superior resistance to penetration by projectiles. 42.7 MACHINERY
42.7.1 PROPULSION SHAFTING
Fig. 42.6 Compositefirewater systems on EXXONs BATT platform. (Courtesy of Ameron Inc.)
are fabricated from pultruded vinyl ester with an outer scrim of continuous glass mat while the pipelines are filament wound glass/epoxy. 42.6.10 CONCRETE FORMS
Repairs to bridges, piers, docks and other structures with deteriorating pilings are accomplished by pulling a fiberglass form around the old piling underwater. The form is set on a precast concrete footing which fits around the old piling. The fiberglass form is then pumped full of concrete. The form may be stripped and re-used or left in place as added protection from ice and floating debris.
Use of composites in propeller shafts is being investigated at the Naval Surface Warfare Center, Carderock Division (Wilhelmi, Appleman and Loo, 1986). The advantages over steel are reduced galvanic corrosion, weight, cost and bearing loads, while allowing an increase in fatigue stresses, flexibility and vibration damping. A small diameter 63.5 mm (2.5 in) filament wound propulsion shaft has been evaluated in a 4-year t i a l aboard a Naval Academy Yard Patrol Vessel (YP-654 Class). Encouraging results from the YP trial, laboratory torsion and fatigue testing and analytical studies employing finite element stress analysis techniques have led to current R&D efforts to develop a Navy-Standard composite shafting 'base laminate', design and procurement specifications and a metal/composite coupling technique suitable for the largest diameter shafting systems.
926 Marine applications
permit measurement of the radial thermal conductivity of experimental candidate composite Composite piping systems are gaining accepheat exchanger tubes (Korczynski, 1993). tance in marine applications (Fig. 42.6). The Installation of composite centrifugal pumps components of piping systems, piping, fittings, valves, heat exchangers and pumps, are (Suitt and Girona, 1993) and piping on Navy all being investigated for replacement with ships has proven encouraging to date. There fiberglass (Wilhelmi, 1988). The major reason are some unresolved issues involved with for replacement is the elimination of corrosion more extensive utilization of composite piping and galvanic effects, resulting in an increased components in Navy ships, namely, concerns service life with less maintenance. In addition, about impact, shock and fire resistance. composite pump components and valves have improved erosion resistance compared with 42.7.3 DIVING EQUIPMENT metallic counterparts. A fluid flow loop was designed and built by NSWC Carderock Composite materials in the diving industry Division to evaluate the performance of the have proven advantages in numerous applicacomposite components under typical operat- tions. Figure 42.7 illustrates a fundamental ing flow rates, pressures and temperatures. shift from the standard Navy Deep Sea Diving This facility also incorporates a single-tube dress consisting of a copper and brass helmet counterflow heat exchanger configuration to and breast plate bolted to a canvas suit with
42.7.2 PIPING SYSTEMS
Fig. 42.7 Navy standard deep sea diving dress, old and new. (Courtesy of Coastal Systems Station, Dahlgren Division, Naval Surface Warfare Center, Panama City, FL.)
Machinery 927 lead shoes and lead weight belt. Out of water the dress weighed over 90kg (200 lb). The modern dress consists of a fiberglass helmet with neoprene suit and weighs less than 45 kg (100 Ib) on land. Many of the buckles and fasteners used to attach lifelines, hot water lines, emergency backpacks and other accessories are made of a thermoplastic with a chopped fiber filler, replacing metal fittings which tend to corrode. Surface supplied gas bottles for the diver’s breathing air are made from fiberglass overwrapped cylinders (Fig. 42.8). These gas bottles can be charged to 20.6 Mpa (3000 psig) and represent a 50% weight reduction over conventional steel cylinders. Figure 42.9 illustrates another type of diving suit. The torso is a carbon fiber reinforced vinyl ester. It is made by hand lay-up on a mandrel and room temperature cured while rotating on a spit. The diver inside the suit remains at one atmosphere of pressure and can dive to 610 m (2000 ft) remaining for several hours without the need for decompression. Portable recompression chambers for transporting divers iA need of treatment from remote areas are made from filament wound Once inside and pressurized, the diver may be transported to a medical recompression facility.
----7
1
-
%
Fig. 42.9 Navy One Man One Atmosphere Diving System (NOMOADS). (Courtesyof Coastal Systems Station, D d g r e n Division, Naval Surface Warfare Center,panamacity,~~.) 42.7.4 CABLES
Because carbon is conductive, some interest has been generated in making light weight conductive cables for underwater instrumentation using specially treated carbon fibers. Due to the weight of copper or aluminum wire, towed electronicsunderwater are difficult to engineer. 42.7.5 BUOYS AND FLOATS
h
Fig. 42.8 Special diver air support system. (Courtesy of Structural Composites Industry, a Harsco Company.)
Sophisticated buoys for oceanographic measurement have been developed for long term mooring of instruments at depths up to 914 m (3000 ft) of sea water. The buoys are made of Syntactic foam and are deployed in a SUbSUrface mode. They can provide over 680kg
928 Marine applications New marine applications of composites are also under development. A particularly challenging and intriguing application is deep diving man-rated pressure hulls. The Advanced Research Projects Agency (ARPA) is sponsoring the Thick Composites Technology Development program to design, fabricate and evaluate a Man-Rated Demonstration Article constructed of carbon reinforced composite materials (Kelly, 1993). The objective of this program is to develop the technology, experience and confidence neces42.8 CURRENT AND FUTURE sary to demonstrate the feasibility of using DEVELOPMENTS composites in primary structure in underwaUse of composites in the marine environment ter vehicles (Hoffman and Kelly, 1992; Leon continues to evolve. There are several recently and Coffenberry, 1992; Camponeschi et al., developed resins, prepreg fabrics, cores and 1994). fabrication processes whose implementation will significantly change the industry. REFERENCES Perhaps the most significant change is the development of effective VARTM processes, Aylor, D. and Murray, J. 1992. The Effect of Seawater Environment on the Galvanic Corrosion most notably SCRIMP (Seemann, 1990).Using Behavior of Graphite/Epoxy Composites conventional resins, fabrics, cores and tooling, Coupled to Metals. Carderock Division, Naval VARTM increases fiber content over hand laySurface Warfare Center SME-92/32. up and almost eliminates VOC emission. Baltek 1986. Baltek Data File 151. Resin development efforts include the for- Burrell, P.P., Herzog, D.J. and McCabe, R.T. 1987. A Study of Permeation Barriers to Prevent Blisters mulation of polyesters and vinyl esters which in Marine Composites and a Novel Technique cure upon exposure to ultraviolet light (Pfund, for Evaluating Blister Formation. Proc. 42nd 1992).The major advantage is that these mateAnnual SPI Conference. rials cure at room temperature but have an Camponeschi, E.T., Jr., Bohlman, R.E., Hall, J. and indefinite working life. Carr, T.T. 1994. The Effect of Assembly Fit-Up The activity in prepreg development is low Gaps on the Compression Response of ThickSection Carbon/Epoxy Composites. ASTM temperature curing epoxies (Pfund, 1993). STP-Compression Response of Composite Materials are available which cure at 85°C Structures, 1994. (185°F) and have several weeks working life, or cure at about 60°C (140°F) with a few days Davies, P. and Lemoine, L. 1992. Nautical Construction with Composite Materials, Proc. working life. Excellent properties can be International Conference Paris, France, 7-9 attained with vacuum bag consolidation December 1992. (Juska, 1993). Garvey, R.E. 1990. Composite Hull for Full-Ocean Depth. MTS Journal, 24(2):49-58. In core technology, low cost thermoplastic honeycombs are being investigated to replace Gougeon 1992. Gougeon Laminating System. Product Literature, Gougeon Brothers lnc. PVC foam and balsa in some applications, Gougeon, M.A. and Bertelsen, W.D. 1993. The particularly to reduce structure-borne noise. Gougeon Hydromat Test System: Special Use of composite core will probably gain Methods and Equipment For Investigating the wider acceptance. In addition, new structural Effects of Various Pressure Loadings on foams are being developed that are not PVCSandwich Composite Panels. Proc. 48th Ann. SPl Conf. and Expo ‘93. based.
(1500 lb) of buoyancy. In addition, buoyancy systems are used on risers for offshore oil rigs. Because drilling takes place in excess of 1.6 km (1mi) deep on some rigs, the drill sections become extremely heavy and require flotation to offset the enormous weight. Syntactic foam with a tough outer composite skin is used for durable flotation. Many Coast Guard buoys are now being deployed with syntactic foam cores and composite skins.
References 929 Greene, E. 1990. Use of Fiberglass Reinforced Plastics in the Marine Industry. Ship Structure Committee Report SSC 360. Hepburn, LCDR R.D., Magliulo, G. and Wright, T.W. 1991. The US Navy's New Coastal Minehunter (MHC): Design, Material and Construction Facilities. Naval Engineers J., May 1991. Hitco 1986. Hitco Technical Report 20-04.16, Rev.C. Hoffman, P. and Kelly, J.J. 1992. Manufacture of Advanced Composite Submarine Structures (MACSS).Proc. 37th Intern. S A M P E Symp. Huss, J. Rowland. 1990. Structural Response of Marine Sandwich Panels to Uniform Pressure Loading. Naval Engineers Thesis, Massachusetts Institute of Technology, Cambridge, MA. Interlux 1993. Technical Bulletin 900B Jacquemet, R. and LaGrange, A. 1988. Aging of Laminated Polyester/Glass and Evolution of Their Mechanical Characteristics in Marine Media. Composites, 4:4046. Johannsen, T.J., 1990. Correct Core Installation, Bonding Techniques and a Repair Case History of a 56 ft Sailboat and a 47 ft Sportfishing Yacht. Paper read at the Atlantic Marine Surveyors Conference on Newest Developments in Boatbuilding and Repairing, 9-10 October 1990, Annapolis, MD. Jones, J.M., Walsh, M., Mansfield, F.B. 1991. Microbial Electrochemical Studies of Coated Steel Exposed to Mixed Microbial Communities. Corrosion 91, Paper 108, NACE. Jones, J.M., Vasanth, K.L., Conrad, T.K., Little, B., Ray, R. 1992a. Corrosion Resistance of Several Conductive Caulks and Sealants From Marine Field Tests and Laboratory Studies With Mixed Communities Containing Sulfate Reducing Bacteria. Intern. Symp. on Microbiologically Influenced Corrosion (MIC) Testing ASTM. Jones, J., Walsh, M., Little, B., Ray, R., Mansfield, F. 1992b. ESEM/EDS Studies of Coated 4140 Steel Exposed to Marine, Mixed Microbial Communities Including SRB. 8th Intern. Congr on Marine Corrosion and Fouling, Taranto, Italy. Juska, T., Mayes, J.S. and Seemann, W. 1993. Mechanical Properties and Impact Damage Resistance of Composites Fabricated by Low Cost, Vacuum Assisted, Resin Transfer Molding. Naval Surface Warfare Center, Carderock Division, SSM-64-93/ 04. Juska, T. 1993. An Evaluation of Low Temperature Curing Prepregs. Naval Surface Warfare Center, Carderock Division, SSM-64-93/07.
Kelly, J.J. 1993. Thick Composites Technology Development Program. Kramers, S.D. Engineering for the Americas Cup. Proc. 1993 Intern. SAMPE Conf., Philadelphia, PA, 25-28 Oct 1993. Korczynski, J.F. 1993. Development of Composite Heat Exchangers For Seawater Applications. Proc. 1993 Intern. SAMPE Conf., Philadelphia, PA, 25-28 Oct 1993. Leon, G.F. and Coffenberry, B. 1992. Proposed Fabrication Process for Thick Composite Submarine Structures. Proc. 37th Intern. S A M P E Symp. Lewit, S. 1990. Use of Polyester Mat as a Bedding Material for Cross-linked PVC Foam Cored Production Laminates. Proc. 3rd Intern. Conj on Marine Applications of Composite Materials, Florida Institute of Technology, Melborne, F1. Little, BJ., Ray, R., Wagner, P., Lewandowski, Z., Lee, W.C., Characklis, W.G., Mansfield, F. 1991. Biofouling, 3:45. Little, B.J. and Wagner, P. 1986.J. Adhesion 20:187. Marino, R., Rockett, T., Rose, V. 1985. Blistering of Glass Reinforced Plastic Marine Materials: A Review. NOAA/Seagrant Technical Report 88, Marine Advisory Service, Narragansett, RI. Miriyala, S.K., Tucker, W.C., Brown, R. and Rockett, T.J. 1993. The Mechanism of Galvanic Blistering in Carbon Fiber Composites. Proc. ICCM/9 Madrid, Spain V.5 pp. 546-553. NRaD 1992. Naval Command, Control and Ocean Surveillance Center Technical Document 2348, San Diego, CA 95152-5000. NSWC, Carderock Division. 1993. Low Cost, High Quality Composite Ship Structures Technology Demonstrated. Research Release, David Taylor Model Basin, Bethesda, Maryland, 20084-5000, May 1993. Owens Corning Fiberglas 1991. 5-2 Glass Armour. Case History. Landing Helicopter Deck (LHD) Deckhouse Armor Application, Owens Corning Fiberglas Corp., Pub. No. 15-ASP-16157-A., Jan 1991. Pfund, 8. 1992. Light-Curing Resins. Professional Boatbuilder Magazine, No. 18, August/ September 1992. Pfund, B. 1993. Pre-preg Reinforcements. Professional Boatbuilder Magazine, No. 24, August/September 1993. Plunkett, J.D., Cohen, R., Kunz, B.P. and Kunz, J. 1992. Light-weight, Volume-Produced Industrial Honeycomb Composites For Marine Structural Applications. MTS '92 Con& Proc.,
930 Marine applications Marine Technology Society, Washington, D.C., October 19-21. Reichard, R.P. 1992. Pressure Testing of FRP Sandwich Panels. M A C M ‘92 Conf. Proc., Composites Education Association, Inc., March 22-24,1992. Russell, M. 1992. Inspection and Repair Manual for Fiber Reinforced Plastic Boats and Craft. NAVSEA Technical Manual T9008-B4-MAN010. Seemann, W. 1990. US Patent Number 4 902 215. Sloan, F. and Nguyen, H. 1992. Applications of Extended-Chain Polyethylene Fibers in the Marine Composites Industry. MACM ’92 Con$ Proc., Composites Education Association, Inc., March 22-24,1992. Smith, C.S. 1990. Design of Marine Structures in Composite Materials, London and New York Elsevier Applied Science. B. 1988. Stachiw, J.D. and Frame, Graphite-Fiber-Reinforced Plastic Pressure Hull Mod 2 for the Advanced Unmanned Search System Vehicle. Naval Ocean Systems Center Technical Report 1245.
Suitt, D. and Girona, F. 1993. Development of a Standard Family of Composite Material Centrifugal Pumps for Naval Surface Ships. Naval Engineers J., May 1993. Tucker, W.C., Brown, R and Russell, L. 1990. Corrosion Between a Graphite/Polymer Composite and Metals. J. Composite Materials 24:92. Tucker, W.C. 1991. Degradation of Graphite/Polymer Composites in Seawater. J. Energy Resources Technol. Trans. ASME 113 4:264268. Tucker, W.C., Little, B.J., Ray, R., Wagner, P. 1993. Microbial Degradation of Fiber Reinforced Polymer Composites. Proc. ICCM/9, Madrid, Spain, V.5, p. 554 Wilhelmi, G.F. 1988. Composites for Ship Machinery Applications. Presented at the SNAME Spring Meeting, 10 June 1988, Pittsburgh, PA. Wilhelmi, G.F., Appleman, W.M. and Loo, T.C. 1986. Composite Shafting for Naval Propulsion Systems. Naval Engineers J., July 1988. Venus-Gusmer 1989.6th Edition Catalog.
COMMERCIAL AND INDUSTRIAL APPLICATIONS OF COMPOSITES
43
Stewart N.Loud
43.1 INTRODUCTION
materials such as E-glass fiber reinforcements, This chapter describes the use of composites and polyester, some of the low-temperature for industrial and commercial applications, performing epoxies, or a variety of engineering which perhaps is best described as thermoplastic resins, usually referred to as 'Miscellaneous' or 'All Other Applications'. It fiberglass-reinforced plastics or FRP (or in is an overview of the myriad number of appli- Europe, GRP). Processes used have been cations that cannot always be neatly grouped mainly the hand lay-up, compression, injecinto major sectors as with transportation, tion, filament winding and pultrusion marine, aerospace/defense, sports, and the methods. Most unit volumes for advanced like. Briefly described are representative appli- composites range from the hundreds of units cations, the benefits offered by composites, per year to a few thousand, unlike the tens of highlights of the primary materials and thousands of units typical with automotive processes, and some conclusions about the parts. Some of the most mature FRP applicaapproximate market size for this extremely tions (yet still with solid growth rates) include diverse market sector. No attempt is made to electrical insulating sheet and sheet molding define the technology of the materials and compound (SMC), bulk molding compound processes, as these are well covered in other insulators, glass fabric reinforced electronic printed circuit boards, pultruded items such as chapters. ladders, filament wound electrical tubing, thermoset SMC and injection molded thermo43.2 HISTORICAL PERSPECTIVE plastic business equipment and appliance housings, hand lay-up antennas, and others. Advanced composites (typically produced This base of FRP applications has been used as using high-performance fibers and/or resins) a launching platform for a huge diversity of historically have been dominant in the aeronew applications. Many now utilize the more space and defense sectors and in many sporting goods products. Hand lay-up with advanced fibers and resin systems or hybrids of these material systems. autoclave molding have been the major processes used for part fabrication. Engineered composites applications for several decades 43.3 PROGRESS have been fabricated predominantly from With the slowdown in worldwide defense spending during the early 1990s, numerous producers have looked to new market opporHandbook of Composites. Edited by S.T. Peters. Published tunities for growth, especially for advanced in 1998by Chapman & Hall, London. ISBN 0 412 54020 7
932 Commercial and industrial applications of composites fiber and resin products, and increasingly to hybrids of several materials. As a result, advanced and hybrid fiber reinforced products are now moving into the application sectors formerly dominated by metals and even glass fiber products. Also, the focus on ’defense conversion’ has resulted in retargeting of some governmental budgets to support defense contractor conversion to civil applications. Many of these applications reside in the commercial and industrial market sectors, the focus of this chapter. A vast diversification of composites has occurred in this sector. From the 1960s and 1970s with many electrical, electronic, appliance and business equipment applications, the business has grown into pressure vessels, laptop computers, centrifuges, flywheels, fuel cells, railroad car components, drive shafts, fuel pumps, loom components, musical instruments, oil and gas riser pipe, power lighting and distribution poles, industrial process rolls, advanced wind turbines and much more. The processes used here have remained fairly constant: pultrusion, filament winding, compression molding, injection molding and hand lay-up, but the use of resin transfer molding has increased dramatically for many of these applications. Materials applications that were almost exclusively fiberglass and polyester, thermoplastics, or low-temperature curing epoxies have evolved into much more vinyl ester usage and migrated to the use of more high-performance fibers such as KevlaP and Twaron@aramids, Spectra@high-molecular weight polyethylene (HMWPE), and many types and forms of the carbon fibers. Ceramic matrix composites (CMCs)and metal matrix composites (MMCs) are of increasing interest and are sometimes competing with polymer-based fiber composites (PMCs), for example in automotive drive shafts. Even carbon fiber reinforced carbon matrix materials (C-C) are moving from the aerospace sector into commercial use. The use of ’advanced’ materials has now matured enough that it is considered by many in the
industry to be counter productive to even use the term ‘advanced composites’ anymore. 43.4 APPLICATIONS REVIEW
Following is an alphabetical review of application segments of this market or of unique and emerging material systems used broadly in related industrial applications. Representative organizations are cited as examples of the types of companies and institutions active in these market sectors so that the reader can research further into the applications of interest. Applications believed at this time to offer the opportunity for significant growth are presented in the most depth. Some applications covered also fit into other market sectors, but their technology is unique or they are industrial components sold into the other markets, so they are described here. Throughout the chapter, tables highlight many applications by major market sector, some long established and many in the emerging growth stage, to stimulate the reader’s thinking on other potential applications that might benefit from applying composites. Those applications in the tables that tend toward more usage of the highest performance materials are listed with an asterisk (*). 43.4.1 ANTENNAS
Telescoping composite tubes are used to form a lightweight antenna support intended for military use and the ultimate in portability. The tubes must be exceptionally light, so the producer used woven Spectra UHMWPE fiber in epoxy resin. They are fabricated using a conventional roll-wrap process similar to that used to produce composite fishing rods. The total weight is only 1.8-2.2 kg (4-5 lb) including 11 nested tubes that collapse to only 711 mm (28 in) for portability but can easily be extended and locked to form a support pole 6.7 mm (22 ft) long. Resin transfer molding is increasingly displacing hand lay-up for fiberglass/polyester
Applicafions review 933 communications antenna dishes. When the economics and performance versus convennew technologies for direct-broadcast satel- tional technologies of steel and concrete. lites become more commercial, small Other organizations are working on a varicomposite dish antennas, probably produced ety of bridge programs where composites are using compression molded SMC, could the central enabling technology for improved become a booming market. civil engineering structures. For example, FRP and CFRP laminated plates are adhesively/mechanically bonded to the 43.4.2 BRIDGES underside of steel or concrete bridge beams to Mostly small foot bridges or the railings, lad- extend their structural in-service life. In ders and decking of structures in highly Scotland, in 1992, an all-composites cablecorrosive environments typically have used stayed footbridge was installed between two FRP components over the last few decades. portions of the Linksleader Golf Club and However, in the last several years there has across the Tay River (Fig. 43.1). Installation been a great increase in interest in the use of was accomplished using mostly on-site hand fiber composites for civil engineering and labor. Much of the structure was constructed infrastructure applications. One segment of - particularly high interest is composite bridges that offer great corrosion resistance and weight savings. Composite tendons or reinforcing bars reinforced with fiberglass, carbon fiber or aramid fiber are being studied aggressively for use as rebar or for prestressing of concrete bridge decks. Defense Advanced Research Projects Agency (DARPA,Arlington, VA) has funded a Technology Reinvestment Project (TRP) at the University of California at San Diego (UCSD) and with the Advanced I Composites Technology Transfer/Bridge )I Renewal Consortium for three technology demonstration areas. First is the development of low-cost fiberglass or carbon fiber reinforced plastic (CFRP) filament wound jackets for the repair and upgrading of freeway bridge or parking structure columns to make them more earthquake resistant. Second is a composite or hybrid composite/concrete bridge deck technology that would allow rapid replacement of bridges on highways and interstates. Last is a plan for a 130 m (450 ft) cable-stayed all-composite vehicular bridge to be built at UCSD over the Interstate freeway at La Tolla, CA, in the late 1990s. All Of the advanced with Fig. 43.1 All-composite cable-stayed bridge, assemfiber, are of high interest for these projects- bled from isopoIyester composites. The bridge’s Also, high-speed, low-cost fabrication A-frame towers are 17.5 m (57 ft) high. (Courtesy of processes will help to achieve competitive Amoco Chemical Co.)
I
934 Commercial and industrial applications of composites Table 43.1 Construction/civilengineering/ infra-
structure Bridge grder upgrade plates* Foot bridges* Cable-stayed bridges* Replacement bridge decks* Earthquake repair and upgrades/column wrap* Bridge decking, panels, enclosures* Pipeline rehabilitation liners* Manhole covers Trench covers Cables, tendons, rebar, dowel bars* Acoustic wall panels* Marine piling Piers* Computer room flooring* Curtain walls*
Glulam wood beams* Mine roof bolts
Earth anchor rods Residential ballistic wall panels Coalescent grid/water treatment* Toxic waste disposal tanks* Chemical grating Platforms
and internal components. For example, Harbec Plastics (Ontario, NY) is a custom injection molder for a compensating arm for a business machine printer housing using 30 v / o fiberglass / polycarbonate compound. They also mold a print head carriage using carbon fiber/PC with Teflon added for wear resistance. Such parts are common to most typewriters and printers. Carbon fiber provides electrostatic dissipation. Equally important is the dimensional stability they obtain to meet a specification for precisionmolded inserts with a 0.025 mm (0.001 in) concentricity on two critical shaft bearings. Growth in this segment comes mainly from further displacement of other materials as composites gain market share. The benefit of high value here is increased design freedom offered by the light and strong FRP materials. This sector is dominated by the molding compounders and the custom molders such as Premix (No. Kingsville, OH) and Rostone (Lafayette, IN).
Walkways Table 43.2 Appliances/businessequipment
from pultruded shapes. In mid-1994, in the UK, a 2720 kg (6000 lb) replacement lift bridge deck for a road across a canal was constructed from pultruded FRP and mounted on steel beams attached to the lift mechanism. Bridge counterweights were reduced, offering cost savings. The infrastructure business sector, particularly roads and bridges, could become one of the largest growth areas for composites to displace steel and concrete over the next decade or more. Development of appropriate codes and standards along with strong cost/performance demonstration projects are keys to success.
Sanitaryware Dishwashers Oven door handles Frying pan components Phone/fax EM1 covers* Copier housings and components* Computer cassette robotic system* Refrigerator liners Business machine bases Laptop computer cases* Microphone booms* Dishwashers Air conditioners Office furniture* Toasters
43.4.3 BUSINESS EQUIPMENT
43.4.4 CABLE CARS
This sector is one of the more mature and has been heavily dominated for many years by compression molded or injection molded reinforced thermoplastics for housings, bases,
Composite cabins were retrofitted to replace steel and aluminum at Chamonix, France to reduce the weight of the cable cars at a ski resort. Ingenex, a French engineering firm,
Applications review 935 built new lightweight car bodies using sandwich construction with Kevlar/epoxy face sheets and a Nomex@honeycomb core. The supporting arm that holds each cable car is fabricated from carbon fiber/epoxy over a Nomex core. These replacement structures allow for many more passengers to be carried on the existing system. This avoided having to scrap the existing system that was becoming overloaded, thus avoiding rebuilding the whole system reportedly at a cost 10 times as much. 43.4.5 CARBON-CARBON COMPOSITES FOR NON-AEROSPACE MARKETS
The vast majority of uses for C-C composites are in aerospace and defense. However, many potential industrial applications are emerging for this unique high-temperature resistant material system. DOE has studied C-C composites for fusion energy systems. One project involved brazing C-C composites. Another program studied development of pitch-based carbon fiber (PBCF) reinforced C-C composite tiles to line the inside walls of future fusion devices such as the Compact Ignition Tokamak (CIT). Perhaps such tiles later will find application in commercial heat exchangers. Kureha Chemical Industry Co., Ltd, and Osaka Gas Co., Ltd of Japan are active in the USA market. Applications targeted include C-C structures for high-temperature furnaces and fuel cells, wet and dry friction materials, plus brake disks and self-lubricating bearings for automotive and truck applications. C-C brakes and clutches are used worldwide in some racing car classes such as Formula 1 and for the Indianapolis 500 race but are banned from other classes due to the premium cost. 43.4.6 CERAMIC-MATRIX COMPOSITES (CMCS)
These materials have received a significant boost in funding because of continuing interest by the USA Department of Energy Office of
Industrial Technologies. Their Continuous Fiber Ceramic Composites (CFCC) Program Plan (10-year plan) was an expansion of the continuing support activity directed toward 'enabling technology' efforts since 1987. DOE defines a CFCC as a long fiber (ceramic, glass or carbon) embedded in a ceramic matrix. Concerns about CMCs raised by potential industrial users include material and processing costs, long-term component fracture toughness, general durability in a variety of environments, the ability to join components with different compositions and the availability of near-net-shape fabrication methods with 3-D fiber architectures (preforms).A relatively new type of ceramics, CFCCs offer the potential to meet the demands for a variety of industrial applications. Some primary applications envisioned for these materials include: burners/combustors, chemical reactors and process equipment, heat recovery systems, refractories and related products, separation/filtration systems, stationary engines, waste incineration systems, tooling, structural components, biomedical parts, wear parts in machinery, and other specialty high-temperature products. Hexcel Corp. (Pleasanton, CA) offers highperformance ceramic materials for various military components such as missile fins, radomes, jet engine parts, and high-temperature tooling, but much larger potential markets might develop in the commercial sector, especially for less-expensive CFCCs using E-glass fiber or other low-cost reinforcements. Examples cited by Hexcel include honeycomb heat exchangers used in high-pressure coalfired gas turbine generators and feed tubes (used to overcome corrosion and thermal shock problems) for a toxic waste incinerator operating at 870°C (1600°F). Successful conversion of the applications noted above could generate savings of 1.1 quads (a quadrillion BTUs) of thermal energy per year, and reduce NOxemissions by 0.6 million tons/year, according to the DOE.
936 Commercial and industrial applications of composifes 43.4.7 CONDUCTIVE FIBER FOR SHIELDING BUSINESS EQUIPMENT
Cytec Industries sells a conductive fiber concentrate or intermediate containing carbon fiber for electrical conductivity. In most of these products, Cytec starts with continuous, standard 33 msi, 12K carbon fiber that is electroplated with nickel before being chopped into 6.35 mm (0.25 in) lengths. The fiber loading is typically 60-75%, even to 90% if sold to a compounder who adds resin to make a specified injection molding compound. 43.4.8 CORROSION-RESISTANT SECONDARY STRUCTURES - GRATING, LADDERS, PIPING
Composite products are offered by many companies including Amalga, Brunswick, Fibergrate, International Grating, Creative Pultrusions and others. Strongwell takes pultruded structural fiberglass shapes (tradename: Extren), fabricates them into useful structures, and distributes to various industrial customers. The material is available in over 100 standard sizes and shapes, and is produced in three series: general-purpose, fireretardant and a premium grade that is both fire retardant and highly corrosion resistant. This grade, called Extren 625, uses a vinyl ester resin; the other two are made with polyester formulations. Applications include handrails, stairways, ladders, cable and channel trays, platforms and walkways, troughs, weir and baffle plates, racks, bridges, identification signs, sump tanks, battery boxes, drain pipes, etc. Besides the offshore drilling and production industry, Strongwell serves a variety of industries, including aquaculture, chemical processing, food/beverage, marine, metal processing, waste treatment, petrochemical, power plants, pulp and paper mills, radar installations, transportation and many specialized applications such as raised computer floors. For static dissipation, carbon fibers can be located at the surface of grating (Fig. 43.2).
I1
Fig. 43.2 ConductiveFRP grating for static electricity dissipation.(Courtesy of CI/SPI.) 43.4.9 D m SHAFTS - PICKUP TRUCKS AND COOLING TOWERS
Strongwell continues to use substantial amounts of carbon fiber for the hybrid drive shafts they sell to Dana for shipment to General Motors. At close to 226 800 kg (500 000 lb) per year of carbon fiber usage and estimated to climb soon to one million lb, this is said to be the largest industrial application for carbon fiber reinforcements. The carbon fibers, mainly from Hexcel, are pultruded with resin along an aluminum tube and overwrapped with fiberglass for damage tolerance. Carbon fiber/epoxy is used for cooling tower drive shafts as well. SpyroTech (Lincoln, NE) is a major supplier of these shafts along with Amalga Composites (W. Allis, WI), Addax, Inc. (Lincoln, NE) and ACPT (Huntington Beach, CA). A circumferential wrap of E-glass/epoxy is added for debulking and impact protection. These shafts exhibit up to five times the durability at one-fifth the weight of steel shafts and are far more corrosion resistant. Cooling tower shafts (as with vehicular shafts) produced with carbon fibers are stiffer and can be produced in lengths twice as long as with steel without suffering vibration damage in use, eliminating the need for intermediate bearings and supports.
43.4.10 ELECTRICALAND ELECTRONICS
insulating hardware, ladders and hotline tools This sector of the market is composed of many for line installers, lighting poles, connectors, applications that are maturing or have high microchip encapsulation and many more. penetration compared with other materials. Fiberglass reinforcement predominates due to Applications include motor and generator its insulating characteristics and low cost. components, switch gear, insulating sheet, Sales are cyclical due to the high market share molded insulators, subway third rail insulat- held by FRP, but there is still strong growth ing cover board, FR-4 and G-10/11 glass fabric seen from penetration gains versus phenolic, reinforced epoxy printed circuit boards porcelain, and other materials. (PCBs), utility transmission and distribution 43.4.11 F
Table 43.3 E l ~ ~ c a l / e l e ~ r o ~ c s / e n e r g ~
systems/commu~cations Electric motors Fuse tubes
Lighting poles Dis~ibutionpoles Cable tray Power tool cases Motor control centers
Printed circuit boards Circuit breakers Guy wire insulators Utility transmission towers Electrical bus bars
Dis~~bution transformer cores, cases Shielding - RFI, ESD, EMI* Battery boxes, casings Electron beam acceIerators* Electrical insulators Electrified third rail cover board Electrical switchgear Power hot-lie tools
Electronic connectors Outlet boxes Aerial lift truck booms Electronics chassis*
Computer chip carriers* Electronic equipment racks" Fuel cells* Flywheel mechanical batteries* Superconducting magnetic energy siorage (SMES) Wind turbine blades*
Telescopingportable antennas* C o ~ ~ i c a t i antenna 5 n ~ dishes Radio masts" Satellite TV dish antmas
Microwave guides Fiber optic cable tension members
~
E
~ BREATHI~G I ~ ~ APPARATUS R
The key component of firefighter breathing apparatus is the air storage bottle that comes in different sizes, materials, and pressure ratings. Structural Composites Industries (Pomona, CA) continues to be well ahead of all competitors making small filament-wound composite pressure vessels. SCI supplies air bottles (all of their firefighter's bottles use S-2 Glass) to Mine Safety Appliances. They reportedly have never had a field failure. Many air breathing systems distributed by MSA are used to escape smoke or other noxious materials in mines, chemical plants, refineries, hotels, aircraft, etc. SCI also is a major supplier for the escape-slide air bottles in use on transport aircraft built by Boeing and Airbus. Many composite oxygen storage bottles are produced for hospital and home care and for use in highflying commercial and general-aviation aircraft (for example, for escape slide inflation but using Kevlar 49/epoxy overwraps for decreased weight). Other composite bottles are used to inflate rafts, balloons and other flotation systems. EFI Corp. (San Jose, CA), a s ~ b s i d j a rof~ Racal Electronics ii-i the UK, and Luxfer USA Ltd (Riverside, CA) continue to be SCI's main competitors as suppliers of firefighter air bottles. They sell air bottles to Scott, Survivair, North Safety and others in the USA and overseas. About 250 fire departments in American cities have bought hundreds of thousands of composite bottles to replace the much heavier steel bottles. Lighter bottles with longer lasting air supplies give firefighters more time on
938 Commercial and industrial applications of composites station in the heat of a fire fight. The bottles must store air for long periods at 31 MPa (4500 psig). 43.4.12 FLYWHEEL MECHANICALBATTERY
SYSTEMS
US Flywheel Systems, Inc. (Camarillo, CA), Calstart (Burbank, CA), American Flywheel Systems (Seattle, WA) and others all hope to supply kinetic energy storage flywheel systems (now often called mechanical storage batteries) for future electrical vehicles in California and elsewhere (Fig. 43.3). Flywheel batteries have an estimated 5-10 times the watt hours per pound storage capacity of conventional lead acid batteries. They are impervious to temperature variations, may be drained repeatedly of stored energy without cyclic degradation, can be recharged in as little as a few minutes, should last 3040 years without requiring replacement or maintenance, and create no toxic waste problems. Mileage range in an electric car (such as GMs Impact) could reach a projected 482-640 km (300400 miles) or more according to some developers, as opposed to a practical range of only 95-128 km (60-80 miles) for today’s electric cars using conventional battery systems.
Fig. 43.3 Flywheel battery components. (Courtesy of American Flywheel Systems.)
Larger flywheels also show promise for stationary energy storage and peak power shaving applications, and the business opportunity for such systems could be enormous. Federal technology centers, such as the Lawrence Livermore and Oak Ridge National Laboratories, now are part of teaming arrangements to enhance flywheel technology transfer from defense and other government programs to the private sector. The rims of the rotating wheels in these systems are filament wound using advanced fiber composites, including E glass, S-2 glass, Kevlar and carbon fibers. At speeds attainable by composite wheels, metal flywheels disintegrate. Use of this system is a good example of technology transfer from aerospace to the civilian sector. An electric car with a flywheel storage system on board (and perhaps with regenerative braking through the flywheels) could approach zero emissions and is expected to meet the much tighter California air quality standards set for early next century. California, New York and Massachusetts, have mandated that 2% of cars sold in those states in 1998 must be zero emission vehicles; the percentages increase to 5% in 2001 and 10% in 2003. The challenge is how to offer an affordable car with desirable range and performance that can meet those goals. Lawrence Livermore National Laboratory reported during a SAMPE seminar that the use of composite flywheels in an electromechanical battery ’would be ideal for use in electric powered vehicles’. They found that a highstrength carbon fiber (such as Toray’s T700 or T1000) yielded the highest energy storage capacity. Since then, Lawrence Livermore National Laboratory and General Motors Corp. have teamed on a three-year, $3 million project to develop a new generation of more efficient automobile flywheel batteries. The GM vehicle project is part of the ’Partnership for New Generation Vehicles program’ launched by President Clinton in September 1993.
Applications review 939 43.4.13 FUEL TANKS FOR NATURAL GAS VEHICLES (NGV)
Lincoln Composites (Lincoln, NE) and ED0 Corp. (Canada) introduced certified 'all-composite' compressed natural gas (CNG) pressure vessels under the 1993 US ANSI/AGA standard called NGV2 for ultralight, nonmetallic natural gas vehicle fuel tanks. The availability of these certified plastic-lined, composite-overwrapped fuel tanks is expected to accelerate the usage of advanced composites into this potentially large alternate-fuel marketplace for composites. Besides weight savings of up to 70% (compared with steel, aluminum, or metal-lined NGV tanks) and greater durability, the new tanks offer increased fuel capacity, extending driving range. The weight savings are achieved through use of a hybrid composite, for example using a 30 msi modulus T650/35 carbon fiber from Amoco Performance Products (Alpharetta, GA) or a Toray T700S/12K carbon fiber. Hybridized with the carbon fiber might be an E-glass fiber roving from Owens-Corning, PPG, or Vetrotex Certainteed to reinforce an epoxy resin formulation. The composite is filament wound over a thermoplastic liner of high-density polyethylene (HDPE). In the hybrid, carbon fiber provides an extremely high strength-to-weight ratio, with the ability to withstand many pressure cycles along with superior fatigue characteristics, while the fiberglass enhances toughness. HDPE liners are lighter weight than steel or aluminum and more resistant than both to the highly corrosive elements sometimes found in natural gas supplies around the USA. A minimum 15-year life expectancy is set for the NGV2 tanks; after that they are to be discarded. Other pressure vessel programs are underway at Comdyne I, Compositek (Kaiser),Fiber Dynamics, Amalga and Structural Composites. Large customers for the older fiberglass/epoxy overwrapped aluminum NGV fuel tanks include General Motors and
Chrysler and many city transit bus operators such as Tacoma, Sacramento, Salt Lake City, Reading, Binghamton and other Northeast and Canadian cities. Recent failures of CNG tanks during refueling (reportedly caused by exposure to unexpected acid solutions) appear to have been solved with an installation redesign and protective coatings. Additional interest comes from utilities such as Southern California Gas & Electric and Columbia Gas companies. Carbon fiber priced today at about $5-6/kg ($ll-l3/lb) for high-strength 12K fiber is about twice as expensive as S-2 glass fiber and much higher than E glass but appears economically competitive for this market. One of the above supplier companies did a cost and weight analysis to decide how much of a premium would have to be paid if higher-strength fibers were used rather than E glass. Using E-glass/epoxy overwrap, the weight would be about 45 kg (100 lb). S-2 glass reduces the weight to 36-38 kg (80-85 lb), and a 12K carbon fiber/epoxy overwrap reduces the weight to 18-20 kg ( 4 0 4 3 lb). Using fiber costs of $2.20, 11-13.00 and 26.50/kg ($1, $5-6 and $12/lb) respectively, the estimated fiber cost per NGV pressure vessel is $64 for E glass, $300-315 for 5-2 glass, and $285-310 for carbon fiber. Combining the various cost components brings the heaviest (E glass) tank cost to $380-400, S-2 tank to $700-720, and the carbon fiber tank to $600-620. Thus, to save about 27.2 kg (60 lb) by switching from E glass to carbon fiber reinforcement, the cost premium is about $220, or $3.6-3.7 per pound saved. Auto designers in the past have suggested that a $2 premium per pound saved was acceptable, but they might be willing to pay more now in view of the environmental benefits gained by burning cleaner fuel. The combined production of mid-size pickup trucks in the USA by Chrysler, Ford and General Motors is reported to be more than 600 000 per year. If only 10% are built with natural gas as the fuel and there are two or three fuel tanks on each truck, there is a
940 Commercial and industrial applications of composites very large potential market in the millions of pounds of materials required. Transit buses may be an even larger market. Table 43.4 Transportation system equipment Seats/frames Monorail cars Mobile storage modules - CNG* CNG/NGV fuel tanks* Rail freight car doors Rail car knuckles Hydrogen storage tanks* Intermodal containers Maglev train guideways* High-speed train brakes* Concept/show car chassis* Electric vehicle frames/chassis* Racing car brakes and clutches* Solar-powered racing vehicles* Highway reflecting markers Airport approach light towers Airport ticket counter tops Drive shafts - pickup trucks* Aircraft escape slide inflation bottles* Ship propulsion shafts* Cablecar cabins and support arms* Shipboard crane components
43.4.14 GLULAM BEAMS
Composites are leading to revolutionary wood beams for bridges; laminated beams benefit dramatically from use of carbon and aramid fibers and cost less. Wood structural building materials have required only the highest quality wood to achieve necessary performance levels. Wood (’the original composite’) has a high strength-to-weight ratio; however, highly engineered components such as beams have had to allow for the inherent variability of wood and wood types. The best woods have come from ’old growth’ forests that today are in decline or are in designated wilderness areas. Logging practices and environmental considerations (remember the spotted owl?) have removed ’old growth’ as a source and have driven producers to consider new technologies.
American Laminators (Drain, OR) and the Wood Science & Technology Institute (Corvalis, OR) have developed a superior line of glued-laminated (glulam) wood beam products based on Douglas fir and composites. Because these beams are a hybrid material system, they marry the light weight and high strength of aramid/carbon fiber composites with the already light weight, high strength nature of wood to make a product that can compete favorably with steel, concrete, and other wood beams. Of special note is that these new composite/wood beams for buildings or bridges outperform conventional wood beam products and usually can be offered at a purchase price or first cost of about 2545% less than conventional glulam beams! Dead weight design load is reduced by 40-60°/0 for large structures and long spans. For structures where deflection is a major consideration in design, the patented technology takes advantage of the high modulus of Akzo Nobel Fortafil carbon fiber with 33 msi modulus and 550 ksi tensile to more than double the design capacity in compression versus typical unreinforced glulam products. Where increased tensile performance is required, Akzo Nobel Fibers’ Twaron or DuPont’s Kevlar 49 aramid yarn is used. Some of these beams can be massive; e.g. typically 24 m long x 1200 mm deep x 254 mm wide (80 ft x 4 ft x 10 in) to much larger. The fibers reinforce pultruded plates that are adhesively laminated during the standard beam forming and bonding. Many structures (100 as of this writing) are already in place using the new composite glulams, in California, Oregon, Wisconsin and Hokkaido, Japan. More structures now are in the design stages (ca. 100) including large domed structures with over 164 m (500 ft) clear spans, for Japan, USA and European clients. Estimates by various participants suggest that the global opportunity for this composite/wood beam technology is in the billions of dollars. In the USA, the conventional
Applications review 941 glulam market is over $1.2 billion/year and is supplied by 147beam manufacturers. Potential requirements in the 45 million kg (100 million lb) range for just the high strength carbon and aramid fibers are seen about the year 2000. One observer close to the project said that 'this has to make a big splash in composites and could be the biggest new composites application in 20 years!' 43.4.15 HIGH-PRESSURE TUBING AND BATTERY CASINGS
Amalga Composites filament winds carbon fiber/epoxy over battery casings for Johnson Controls. The Inconel liner has a carbon fiber/epoxy overwrap which is under 0.34 MPa (50 psi) stress when discharged and 5.86 MPa (850 psi) when fully charged. These batteries are designed to be buried in the desert to power instruments during the night and they are recharged during the day. Amalga also makes vessels for high-pressure applications, including a filament wound carbon fiber/epoxy-overwrapped steel tank to withstand 3.41 MPa (500psi) for use as an emergency backup pressure cell for the F-22 fighter. Other interesting high-pressure products are carbon fiber/epoxy-reinforced metallic cylinders that achieve 41-83 MPa (6000-12 000 psi) of pressure capability. Lower pressure pneumatic tubes are being made by Amalga for service at 1.7 MPa (250 psi), and these usually are reinforced with a fiberglass/epoxy overwrap. 43.4.16 HIGH-SPEED TRAIN BRAKES CARBON-CARBON
Carbone Industrie in France has been testing carbon-carbon brakes on the French TGV Atlantique line. Two similar but slightly different applications are being tested; for the drive wheels and for the bogey. The C-C brakes must qualify on cars being stopped from 300-350 km/h (to 215 mph) before they can be considered for regular service. Alsthom, the
TGV car designer and builder, must decide whether the weight savings, lack of fade and projected longer life (as with similar C-C aircraft brakes systems) justify a change in the steel brakes now in service. With plans to build similar high-speed rail systems in Texas and other parts of the USA, there is growing interest in the results of the tests. Several companies in Japan have similar interest in using C-C brakes on proposed advanced bullet trains, which also might reach speeds of 350 km/h. Later magnetic levitation (maglev) trains which will approach 500 km/h may be assessed for use of these ultra-high-performance brakes. 43.4.17 HYDROGEN FUEL STORAGE
Hydrogen storage tanks may offer potential for filament-wound pressure vessels, but it will be several years before this can possibly occur. Syracuse University received a grant from the Department of Energy to prove a novel approach to hydrogen storage. The hydrogen would be stored in a filamentwound tank in which activated carbon granules are closely packed. After being cooled in a liquid nitrogen heat exchanger to the desired storage temperature, the hydrogen is passed through the activated carbon where some is adsorbed. It is then stored at low temperature (150K) and relatively low pressure 5.51 MPa (800 psig) before being released via slight heating of the granular carbon while opening a valve for controlled release into the nearby hydrogen-oxygen fuel cell. A design engineer predicts that a car powered by an electric motor energized with a hydrogen-oxygen fuel cell will be able to outperform a similar vehicle powered by a gasoline engine. Since hydrogen as a fuel requires a massive change to the fueling infrastructure, the timing for this market sector is hard to predict. If such an infrastructure is installed, there will likely emerge many more composites applications such as large tanks and pipelines.
942 Commercial and industrial applications of conqposites Table 43.5 Miscellaneous ~~
"JurassicPark dinosaurs High-altitude balloon gondolas* Aquaculture cages for fish farming Telescope tubes* Helmets* Superconductingsupercollider* Firefighter air bottles* Oxygen storage bottles* Life raft inflation bottles* Flag poles Deep-submergencecapsules* Riot control body armor/shields* 43.4.18 LAPTOP COMPUTERS
Lightweight laptop computers with carbon fiber reinforced plastic cases were first introduced in the USA in the early 1990s at the giant Comdex computer show in Las Vegas. Toshiba America Information Systems (Irvine, CA) introduced three models, reporting that their light weight 2.5-3.6 kg (5.5-7.9 lb) and portability were partially due to the use of a carbon fiber molded case (Fig. 43.4). The new cases
molded with nonwoven-fabric are 20-50% lighter than older, thicker plastic cases and have a wall thickness of only 3.04 mm (0.12 in) without sacrificing ruggedness. The carbon fiber/thermoplastic is supplied by Kobe Steel Ltd. and Mitsubishi Rayon Co., Ltd of Japan. In a related competitive development, Advanced Logic Research, Inc. (ALR - Irvine, CA), introduced the world's first modular notebook PCs offering processor chip upgrades. The ALR notebooks are ergonomically designed with a thin carbon fiber reinforced, lightweight case that fits into a standard-size briefcase. Other notebook computer producers have composite development programs underway. Based on market share estimates provided by Stanford Resources, Inc. in PC Magazine (August 1994), Compaq has the largest market share today at 14.1% of shipments in 1993, Apple is second at 13.6?'0, Toshiba is third at 11.1%, IBM is fourth at 7.1%, followed by dozens of other producers. 43.4.19 LOOM COMPONENTS
DFI Pultruded Composites, Inc. (Erlanger, KY) was one of the first to successfully pultrude carbon fiber/epoxy, once quick-curing epoxies of the proper viscosity became available. One of their better customers is in Europe for DFI's pultruded lightweight, low inertia, carbon fiber shuttle arms that rapidly reciprocate in the older style but updated rapier looms still being used by many industrial fabric weavers. The shuttle arms are delivered in six lengths ranging from approximately 1.3 to 2.1 m. DFI used 34 msi modulus carbon fibers as the reinforcement. [Incidently, one of their unique products is a pultruded rod using carbon fiber/epoxy with a diameter as small as 0.25 mm (0.010 in).] Polygon Company Div. of Plas/Steel Industries (Walkerton, IN) pultrudes a carbon fiber/epoxy, high-speed loom frame compoFig. 43.4 Lightweight computer case with nonwoven fiber reinforcement. (Courtesy of Toshiba nent for a USA equipment producer, using America InformationSystems, Inc.) Akzo Nobel Fortafil's 50K filament fiber.
Applications review 943 43.4.20 MAGLEV TRAIN GUIDEWAYS
Magnetic levitation or maglev train system concept definition contracts are funded by the DOT-Federal Railway Administration under the $151 billion Intermodal Surface Transportation Act which was passed in 1991 and included $750 million to be used specifically to complete the design and construction of a USA demonstration maglev line about 32 km (20 mile) long. The German Transrapid maglev system design is planned for the 20km (13 mile) demonstration system to be built in Orlando, FL. USA and Canadian companies competing for future maglev system contracts include Northrop Grumman, Foster-Miller, Magneplane, Boeing, Bombardier and Morrison-Knudsen (the latter three are proposing the most extensive use of composites on the ‘rolling’ stock, perhaps including the outer skins of the maglev cars plus the interiors). On the maglev vehicles, General Electric is developing shielded superconducting magnets and their coil design, shielding, motors, levitation and carbon fiber/epoxy cryostats. To employ the latest methods of producing low-cost composite train vehicle structures and to keep costs down, carbon fiber will be used selectively, with high-performance fiberglass the main reinforcement for sandwich panels, possible low-FST (fire, smoke toxicity) materials for the vehicle interiors. Regarding the guideway, they favor a Ushaped configuration and either reinforced concrete throughout or with a composite grille at the center of each guideway beam. To reinforce the concrete, they recommend using post-tensioned FRP tendons in the magnetic field and closest to the path of the cars and more conventional steel rebar prestressing in the lower regions of the supporting columns, well away from the high magnetic field to avoid raising the maglev drag coefficient significantly. Magneplane International predicts that the guideway will be the dominant cost component, so it has first design priority. Their target is not to exceed $6.2 million per
km ($10 million per mile) for a dual track guideway. Their design includes a concrete box beam for the guideway with the bottom of the vehicle ’wrapped’ around the beam for safety. The guideway might use a hybrid of fiberglass and carbon fiber rebar rather than steel rebar reinforcement. Foster-Miller is focused on high-performance, low-cost guideways. Most of their current emphasis is on using high-strength FRP post-tensioning members (both transverse and longitudinal) in the upper areas of the guideway. West Virginia University is working on a DOT-FRA contract titled: ’Assessment of Maglev Guideway Systems’. Imbedded fiber optic systems may be used to monitor entire guideway systems and to detect potential trouble spots such as joints. West Virginia University data show that if FRP rebar or tendons are used, the load-carrying capacity is about twice that achieved with steel. Although the initial cost might be five times greater, the weight reduction can be 80%, so on a performance-per-pound basis, FRP would be lower in installed cost than steel rebar reinforced concrete. WVU also is looking at more advanced composite materials, including carbon fiber and hybrids. Another option is to use an aramid fiber tendon, and candidates include Teijin’s Technora, Akzo Nobel Fiber’s Twaron and Du Pont’s Kevlar. 43.4.21 MEDICAL, HEALTH AND SAFETY PRODUCTS (SEE CHAPTER 44)
BioMechanical Materials Corp. (Camarillo, CA) and Composiflex (Erie, PA) are among those selling carbon fiber/acrylic and carbon fiber/urethane sheet stock to orthotic equipment distributors. They in turn sell this material to hundreds of small local laboratories that fabricate components tailored by orthotists for each patient, including such items as arch supports, knee braces and custom sports braces. BioMechanical Materials among others also sells large quantities of composite sheet
944 Commercia2 and industrial app2ications of composites and fabricated parts to sports shoe manufacturers, such as Brooks and L.A. Gear, for enhancing performance and the marketability of the top-end shoes. Hexcel supplies a unique hybrid E-glass and carbon fiber fabric reinforced epoxy 'chip' for arch supports in Reebok sports shoes. For electrical hazard protection, Composiflex produces a hybrid composite of fiberglass/carbon fiber/epoxy to fabricate lightweight toe plates for safety boots and puncture-resistant insoles, replacing the traditional steel inserts. This company gets its name from their product with multiple plies of oriented fiber composite which can control bending to certain directions. Table 43.6 Medical products MRI scanner cryogenic tubes* X-ray tables* Underwater therapy treadmill Wheelchairs* Orthotics - shoe arch, neck & knee braces* Prostheses - artificial limbs* External fixaters - orthopaedics* Rehabilitation baths Medical centrifuges* Hip implants* Heart valves* Dental/prosthetic cements X-ray file cassettes* Medical sensors* Walking canes* Artificial ligaments* 43.4.22 METAL-MATRIX COMPOSITES (MMCS)
Duralcan USA (Novi, MI) markets low-cost MMCs reinforced with alumina or silicon carbide particulate reinforcement. One of their most successful applications is the use of reinforced aluminum wire for tire studs. Finland and Sweden (and many USA states) banned the older steel-jacketed tungsten carbide studs from all-weather/snow tires due to road wear. Many people assumed that studs degrade highways prematurely because they wear away road surface. The real culprit is centrifugal force acting on the studs that causes the
older heavy tungsten studs to impact (rather than scrape) the road surface. A Duralcan MMC with 15 v/o alumina in 6061 aluminum is being substituted to limit road damage because of its lower weight. MMC wire is being shipped at the rate of 'several tons per month' for this application. Another interesting wire application is for flame spraying where the MMC adds to service life; aluminareinforced aluminum wire is sprayed onto steel as an antiskid surface in marine environments such as ships, piers and oil platforms. Previously, pure aluminum was flame sprayed onto a steel surface before painting, serving as a corrosion barrier. Among the many other applications in work or which have been tested to date are the following: extruded truck frame channels [Cchannels of about 304 x 76 x 7620 mm (12 x 3 x 25 in)], motorcycle drive sprockets, a proportional valve for a truck lift gate, compressor piston-ring inserts, sporting goods such as golf clubs and brake components for standard and high-speed rail cars. Other promising applications are automobile brake rotors (initial orders are now being used for early production at this time) to reduce weight by replacing cast iron. Further component applications likely to emerge include pickup truck drive shafts, engine cylinder liners, automotive suspension arms, aircraft camera mounts, electrical housings, and pistons for automotive engines, piston inserts that reduce the energy required for high-speed reciprocating motion in large gas compressors, and tubing and forks for high-performance racing and mountain bicycles. Recyclability is a major market issue for high-tonnage applications of many materials, especially now in Europe. Duralcan claims that their MMCs have a potential advantage over many other materials, including polymer-matrix composites, since MMCs can be recycled as a composite material or easily reclaimed into high-quality aluminum by melting and removing the reinforcing particles. Whether recycled or reclaimed, Duralcan
Applications review 945 USA can achieve virgin material properties in new parts. For example, they have recycled pickup truck drive shaft extrusions up to four times, while retaining original property performance in subsequent parts. In the aluminum casting process, 30-40% of the process metal is not in the part but in the gates and risers, and all of this material is now being recycled. Duralcan USA's parent company is Alcan Aluminium Ltd in Canada (Jonquiere, Quebec), and their plant has an ultimate capacity for these products of around 11-16 million kg (25-36 million Ib) per year. In some applications (bikes and drive shafts, for example) these MMCs could readily displace some growth opportunities now seen for polymer composites. MMCs from other producers also show great promise in the field of electronics. Multichip modules (MCMs) require high dimensional stability, thermal conductivity and thermal shock resistance. Mounting them on MMC provides these benefits, especially for high-density avionics applications on aircraft or for radar and other systems on board ships. For maximum thermal conductivity in electronic backplanes, Amoco Performance Products's K-1100 pitch-based carbon fiber might be used.
Composite trailers could haul more fuel or would be allowed on roads with weight restrictions or service pads not capable of handling the heavier steel trailers. Manufacturers of the current steel tube trailers include Christi Park Industries in McKeesport (PA) and Trend Fuels in Austin (TX). Certification by the Department of Transportation (DOT) is required to allow use of large composite storage tanks on highways if loaded with compressed natural gas to 25 MPa (3600 psi) or higher. Projections for the composites potential in this one application range from 3645 million kg (80-100 million lb). 43.4.24 MOLDING COMPOUNDS FOR INJECTION MOLDING
A steadily growing segment of the advanced composites market is a variety of injection molded parts made with carbon fiber, or hybrids with fiberglass or aramid reinforcements. LNP Engineering Plastics, Inc. (a unit of Kawasaki Steel based in Exton, PA) participates in many current applications in the commercial and industrial sectors. (See Section 43.4.3.) There is a trend toward increased use of the higher-temperature thermoplastics, and at LNP these include polyethersulfone (PES), polyarylsulfone (PAS), polyetheretherketone (PEEK), polyphenylene sulfide (PPS) and polyetherim43.4.23 MOBILE STORAGE MODULES FOR CNG ide (PEI, or Ultem). Their long fiber Aerojet and Pacific Gas & Electric have teamed compounds called Verton offer performance at to demonstrate a low-cost, lightweight com- the high end of the mechanical performance posite Mobile Storage Module that is a large spectrum for compounds. RTP Company (Winona, MN), Comalloy NGV fuel tank made of carbon fiber or fiber(Nashville, TN) and glass composite overwrapped aluminum used International to transport CNG from the utility to the user. Compounding Technology, Inc. (Corona, CA) These tubes are currently 381 mm (15 in) manufacture molding compounds made with diameter by 6096 mm (20 ft) long, and several the high-performance thermoplastic resins of these would be mounted in a truck trailer listed above plus large amounts of polycarthat can be used to transport, store, and bonates and nylons/polyamides, much of deliver natural gas fuel to fixed loading areas. these for electrical and electronic insulating These modules would be moved between fleet components. A smaller part of the business is refueling stations, thereby avoiding the neces- nickel-coated carbon fiber for added conducsity of installing a costly pumping system. tivity, and they also use some stainless steel
946 Commercial and industrial applications of composites fiber to make compounds suitable for EMI/RFI Shielding. DSM Engineering Plastics, Inc. (Evansville, IN) and others report a noticeable trend toward using higher-temperature glass-filled thermoplastics for components used in the electronics industry, mainly for connectors. Currently, the liquid crystal polymers (mainly Vectra from Hoechst Celanese and Xydar from Amoco) are widely used for electronics applications. Having been part of the Akzo organization for several years (and before that, Fiberfil), they have worked closely with Akzo Nobel Fortafil Fibers (Rockwood, TN) to develop special grades of chopped and milled carbon fibers with suitable finishes. Other suppliers include Toho Carbon Fibers and Hexel. Heavy-tow fiber (50K filament count and greater) made from commercial-grade PAN precursor is well suited to making lower-cost chopped fiber. Usage of carbon fiber has been going up every year and has reportedly passed annual consumption of 2.25 million kg (5 million lb) per year that would translate into compound volumes five to ten times that. 43.4.25 MUSICAL INSTRUMENTS
Kaman Aerospace (Bloomfield, CT) offers the Ovation Instruments guitar. The top of the Ovation series is called the Adamas and its excellent tone is attributed to the use of a composite sandwich construction for the top. The outer and inner plies are single layers of 5-mil unidirectional carbon fiber/epoxy tape, while the core is 0.762 mm (0.030 in) thick birch veneer. Fiberglass composites are also used for other Adamas components, including the bowl. These are mostly handmade guitars so annual production is small. A producer of classical guitars is Kuau Technology (Maui, Hawaii) with carbon fiber/epoxy and the RTM process is their preference. The Luthier’s Mercantile (Healdsburg, CA) catalog for stringed musical instrument producers generates 99% of their business via mail order. They offer carbon fiber/epoxy bar and sheet stock in
various sizes for reinforcing the necks of guitars or banjos, plus for making bridge blanks, fingerboard underlays, stiffening braces and other components requiring extra strength, stiffness and/or creep resistance. Other instruments now offered in carbon fiber and other composite materials, all claiming better tone and clarity, are violins and bows, bases, drums and sticks, sousaphones/tubas (an old application in FRP), harps, pianos and more. Table 43.7 Consumer products
Camera tripods* SCUBA tanks* Banjos* Sousaphones/ tubas Harps* Pianos Umbrellas* Attache/brief cases* Cassette tape/CD player cases* Exercise equipment* Art work bracing systems* Guitars*
Violins and bows* Bases* Camera cases*
Eyeglass frames* Caskets
Binocular /monocular bodies* Running shoe arches* Audio earphones* Drums and sticks*
43.4.26 OIL AND GAS APPLICATIONS
Conoco, Inc. (Ponca City, OK), Shell (Houston,
TX) and others are involved in oil and gas exploration to production depths typically to 619-928 m (2000-3000 ft) and to even more than 4 km (15 000 ft). Conoco is the first company with one of the new tension-leg platforms (TLP) in use. Shell is building such a platform and planning several more in partnership with Exxon and British Petroleum. The oil industry has a ’steel and concrete bias’ which is the result
Applications review 947 of many years of experience, relatively low initial cost, ample supply, current design codes, etc. However, new materials are the major technology thrust needed in offshore structures because the industry is forced to spend too much money fighting corrosion of steel. Current offshore structural applications of composites include low-pressure pipe, gratings, handrails, equipment covers or enclosures and ladders. Future application development is likely to focus on FRP for high-pressure piping, pressure vessels, accommodation modules and blast walls. FRP products (mainly E-glass reinforced) provide installation savings ranging from 10-70%, improved life-cycle costs, enhanced safety, 30-70% weight savings and environmental advantages, according to studies by Conoco. It is expected that a variety of composite materials will be used for offshore structures, with carbon fiber, aramid and fiberglass reinforcements, thermosets and thermoplastic matrix resins and some hybrids. The Institute Franfais du Petrole supported development of advanced composite tubes developed by Aerospatiale for risers on offshore stationary drilling platforms and tension leg platforms. These tubes are hybrids of highstrength fiberglass and carbon fiber and are offered in diameters ranging from 76 to 300 mm (3 to 12 in) and in lengths up to 24 m (80 ft). Internal working pressures are in the 34-103 MPa (5000-15 000 psi) range, tensile loads to 90 000 kg (100 tons) will be accommodated by a special high-strength threaded steel alloy coupling, and the design life can be as high as 20 years. Typical drilling risers will be up to 10 times the height of the Eiffel Tower [up to 3000 m (9800 ft)] and weigh over 2.2 x lo6 kg (2500 tons). Syntactic foams control buoyancy and reduce the effects of sea loads from waves, currents and storms. The case for composites is clear, as weight savings could exceed 544 000 kg (600 tons). A 900 kg (one ton) reduction in the offshore superstructure or the risers results in a savings of 272 kg (3 tons) of displacement load on the platform. Working
closely with IFP and Aerospatiale Space & Strategic Systems are Lincoln Composites (Lincoln, NE and teamed with Aarding BV in The Netherlands) plus Coflexip & Services, Inc. (Paris). Coflexip is reportedly the leading producer of sub-sea pipelines and risers to transport oil and gas products at offshore production facilities. The initial product of the Lincoln-French joint venture was a composite TLP riser pipe that connects well-heads to tethered floating platforms. Risers provide a casing to protect the well bore and to carry the oil to the surface. Tethers will be 244.5mm (9.60 in) diameter tubes, with a wall thickness of about 13.5mm (0.53 in) and 12-17m (40-55 ft) long. Table 43.8 Oil and gas exploration
Tension leg platform tethers* Fire shields - drilling rigs Drilling tubes* Sucker rods/slides Oil well logging probes* Tribological bushings, seals, ' 0 rings* Drilling motor shaft Drill casing
Longer-term commercial application growth in this sector for advanced composites should include moorings and tendons, drilling and production risers, drill pipe and tubing and some pressure vessels. With platform construction costs exceeding $1.5 billion per unit, there could be a substantial dollar potential for the various composite applications. One producer estimates the current usage of steel to be as high as 1.5-1.8 x lo9kg (1.7-2.0 million tons) used for oil- and gas-well downhole tubulars. Much of that could be displaced by composites (at one fifth the weight) because of their advantages in high specific strength, fatigue resistance, corrosion resistance, lighter weight, low coefficient of thermal expansion and reduced installation cost (due to lighter weight)
948 Commercial and industrial applications of composites 43.4.27 PIPELINE REHABILITATION
A high-volume application for Kevlar 49 and other aramid fiber reinforcement has been developed by Insituform of North America, Inc. (Memphis, TN) with the help of Du Pont. A patented product called Insituform has been used extensively for years in repairing and generally 'rehabilitating' old large-diameter, gravity-flow and pressure pipes. Insituform can repair or replace a damaged pipe very rapidly to minimize downtime. The process was developed in the UK many years ago but most of the recent progress has reportedly been made in the USA. A special needled polyester felt reconstruction tube (called Insitutube) is coated on the outside (typically with a thin layer of polyurethane), and is custom engineered and manufactured to fit the damaged pipe exactly. It is impregnated with a liquid thermosetting resin (polyester, vinyl ester, or epoxy as appropriate for the chemical environment to be experienced) and lowered into a manhole through an inversion tube. One end of the Insitutube is attached to the lower end of the inversion tube elbow. The resinimpregnated tube is turned inside out using cool water so that the smooth coated felt exterior then becomes the new interior surface of the pipe. Circulation of hot water through the tube then cures the resin-impregnated felt. This repair process typically is performed on an old water or sewage pipe that has been buried for many years and which is starting to leak. The whole procedure can be done without digging up or otherwise disrupting the old pipe. Kevlar 49 is used to reinforce the polyester felt in the hoop direction. This reduces the felt thickness and the total weight, increasing the inside diameter of the repaired pipe, and generally adds both structural strength and rigidity. Also involved is Shell Chemical (Houston, TX) where a special quick-curing epoxy is being developed for this application. Dow Plastics (Midland, MI) supplies resins to another company in this business, In-Liner Corp., with their cured-in-place-pipe (CIPP) process, among many users.
43.4.28 POLES - POWER DISTRIBUTION AND
LIGHTING Shakespeare (Newberry, SC) developed filament wound distribution poles that were installed by Montana Power Co. These fiberglass/polyester poles weigh half as much as wood; material efficiency is improved through the use of stitched fabric reinforcements as well. Weight savings enhance the ease of installation in remote areas through use of smaller helicopters than required for wood poles (Fig. 43.5). The 355 mm (14 in) diameter poles are generally 12-14 m (40-45 ft) tall. Another major advantage of FRP poles is the elimination of chemical treatments required to preserve wood poles. An 80-year service life is
. ... .
'
.,
.4.
Fig. 43.5 Installing lightweight composite utility poles by helicopter. (Courtesy of Amoco Chemical CO.)
Applications review 949 targeted. In a related development, Composite Power Co. (Las Vegas, NV)fabricates FRP utility poles using pultrusion. Prospects for both companies are enhanced because of the environmental concerns when using treated wood poles and the cost escalation of older products that bring them closer to the current premium costs of composite poles. The California DOT (Sacramento, CA) is potentially interested in the use of composites to build 53 m (175 ft) tall light standards for highway interchanges. Caltrans currently uses galvanized steel but these must be replaced periodically and are difficult to erect and maintain. Among the firms that have built tall steel light standards are Ameron, Union Steel, Valmont and Shakespeare. The hope is that current composite technology for lighting and power poles and for sailboat spars to 55 m (180 ft) height can be expanded into such highway infrastructure applications.
fiberglass knitted multiaxial fabrics from Brunswick Technologies, Inc. (Brunswick,ME) and vinyl ester resin from Dow Plastics (Midland, MI). The process is called Seemann Composite Resin Infusion Molding System or SCRIMP. Co-owners of the process technology that can infuse preforms of 6 m by 2 m (20 ft by 7 ft) in only 35 mm (1.4 in) are Hardcore, Seemann Fiberglass and TPI (also see rail item above). They license this process technology to others. The USA Navy has been assessing this technology for years in the development of major ship structures, including masts. Composites sheet pile from companies such as Creative Pultrusions and others are also likely to experience a major growth trend. 43.4.30 RAILROAD ROLLING STOCK COMPONENTS
Burlington Northern Santa Fe Railroad (Ft. Worth, TX) is a very innovative company and plans use of composite overwrapped air bot43.4.29 PORTS AND HARBORS tles for use on locomotives. They favor the A potentially significant new application with lightest version that would use a plastic liner major potential for tonnage of FRP is that of and a carbon fiber/epoxy overwrap to replace dock and wharf pile fenders. Du Pont fabri- heavy steel tanks for improved corrosion resiscated composite fenders for a systems tance and easier/safer installation by work company. Six developmental fenders were crews. Weight saved on railroad rolling stock installed in mid-1994 in the Chesapeake Bay also increases the capacity for revenue generaarea. They replaced multiple sets of fender pil- tion from hauling increased cargo. Lincoln ings in 10-pile clusters that had steel fender Composites and Compositek fabricate such air surfaces. The treated wood pilings only last bottles while Dresser Industries’ Nil-Cor about seven years on average, the steel plates Operations makes a variety of composite are drastically deformed and rust relatively valves. Most use fiberglass/vinyl ester construction; others are made of carbon quickly. The FRP fenders with UHMW polyethylene fiber/vinyl ester, fiberglass/polysulfone and skinned impact surfaces to date have better fiberglass/polyphenylene sulfide composite. BNSF also seeks to introduce composites on withstood the impact of a 159 000 kg (350 000 lb) ferry boat (eight to ten times/day) existing or new freight cars. A prototype comthan did the wood/steel system, with no dam- posite knuckle is in the design stage for age reported in over two months as of this extended field tests. This is likely to be a writing. The lightweight composite fenders hybrid featuring carbon fiber/epoxy with were installed in only 3.5 h with a cherry steel inserts to protect the high-wear areas. picker while the steel and wood system Another industry program is development of require a barge crane at much greater expense. an FRP cargo door to be produced using a Composite fenders are fabricated using resin infusion process called SCRIMP offered
950 Commercial and industrial applications of composites by Hardcore Composites (New Castle, DE), Seemann Fiberglass (Hanrahan, LA), and TPI (Warren, RI). Composite auto hauler composite cars with all-pultruded components of fiberglass/polyester have been developed by W. Brandt Goldsworthy & Assocs. (Torrance, CA) along with Alcoa Composites and Stoughton Composites (Brodhead, WI) for the Union Pacific. Many other freight car development programs are under way at Trinity Industries and elsewhere. Corrosion resistance and light weight are the two primary benefits of composites that are sparking this interest. 43.4.31 RAILWAYS AND TUNNELS
Besides pultruded and filament wound composite cableways, ducts and piping, there is a vast array of associated molded parts for trackside, station equipment and furniture. Development of fiberglass reinforced phenolic-matrix composites, with their superior FST properties, has become an important contributor to the increasing acceptance of composites for interior components. European countries now virtually mandate using phenolic resin composites for underground railcar and station components. Some of ICI’s Modar acrylic-modified polyester and also the less flammable phenolic-based composites were used in the Channel Tunnel, mainly for electrical cable tray. The market for composites in railways and tunnels could double during the next few years due mainly to the greater attention being paid to safety, including improved fire, smoke, and toxicity performance (reduced FST). 43.4.32 ROLLS AND AIR SHAFTS
Industrial rolls and expanding air shafts used inside cardboard core tubes on winders to take up paper, films and foils will be major growth applications in the industrial composites sector. Tidland Corp. is said to be the world’s leading supplier of core shafts of all types and
reportedly buy most of their carbon fiber/epoxy rolls from Lincoln Composites. Shafts from about 2-10 m (6-33 ft) long are fabricated using hybrid carbon fiber and fiberglass in epoxy which are about one-fifth the weight, operate at faster speeds and use less energy than the conventional steel rolls. Fiberite (Wilmington, DE), Epoch Industries (Garland, TX), Addax (Lincoln, NE), and Quality Composite, Inc. (Sandy, UT) are among others that fabricate such products. Amalga Composites, Inc. (Milwaukee, WI) is a more recent entry into the filament winding of carbon fiber/epoxy tubes for the industrial rolls used in paper and plastic film converting machinery and reportedly is a major supplier to American Roller (Bannockburn, IL). They use Grafil’s 33 msi carbon fiber but leave the finishing and coating operations to American Roller. According to ARC, there are many potential applications for composite rolls. These include replacing the steel rolls commonly used in winder core, winder shaft, idler, layon and dancer/ transducer positions to manufacture film, flexible packaging, foils, magnetic audio and video tape and photoand pressure-sensitive papers on printing presses, laminators, coaters and blown film lines. However, selecting the right composite for each of these positions is not quite as simple as ordering an off-the-shelf product. ARC technical and marketing people make a special point of explaining why each composite roll should be customized to the end-user’s specific application and operating parameters, especially because of the many types of composite materials and structural designs available on the market today. Another filament wound tube fabricator for this application is SpyroTech Corp. (Lincoln, NE). Also, TPI (Warren, RI, and formerly called Tillotson Pearson) has been using filament winding to make fiberglass/epoxy and carbon fiber/epoxy tubes for paper mill rolls. It is noteworthy that Mitsubishi Rayon in Japan is one of the largest suppliers of such rolls in the world. These suppliers market the
Applications review 951 major benefit of hybrid carbon fiber-fiberglass and 120°C (250°F) cure epoxy composite rolls and shafts compared to steel and aluminum shafts, namely a 50% weight reduction. Precision balancing allows higher operating speeds than with metal rolls. Advanced composite rolls exhibit 20% less deflection under loads. Also, composite rolls offer superior fatigue resistance for increased service life and improved damage tolerance versus steel or aluminum. Polygon (Walkerton, IN) is a pultruder and braider active in fabrication of such industrial rolls. Composite Development Corp. (West Wareham, MA) also supplies to the converting industry that wants to use thinner gage paper/film/foils yet process at higher speeds, so the lightweight, low-inertia rolls produced using carbon fiber composites with the fabric roll-wrap process are rolls that offer improved handling characteristics (Fig. 43.6). Heavier rolls such as plate cylinder rolls (used to mount printing plates) offer excellent weight savings that are important in a process with constant and many job changeovers. Some producers of rolls use braided reinforcement preforms from companies such as A&P Technology (Covington, KY).
'
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I ~
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Fig. 43.6 Composite industrial rolls. (Courtesy of Composite Development Corporation.)
Table 43.9 Industrial equipment
Coordinate measurement/3-D gaging* Magnet support rings* Torque tubes* Paper folding plates* Coating doctor blades* Textile bobbins* Instrument housings Pump impellers* Crane components Backhoe components Forklift components Mixer drive shafts* Containers Chemical mixer paddles* Liquid level measuring sticks Surveying transit tipod Industrial rolls - paper, film,foil, printing* Water tanks Chemical filtedframes Radio-controlledvehicles Mirrors* Cryogenic vessels* Core tube air shafts* Generator housings Welding robot arms* Laser resonator spacer tubes* Pump casings Speaker housings* Signage Textile looms - rapiers, heddles, shuttle cocks, picker sticks* Ball and butterflv valves Conveyor belt sGpport shafts High-pressure hydraulic pipe* Pneumatic piping' Uranium enrichment centrifuges* Cooling tower drive shafts* Tote trays/Storage bins Tooling* Optical benches* PCB conveyor components Coil springs* Underwater light housings Light bulb sockets Propellers Fan blades Ladders Safety shoe toe guards Industrial shaft couplings Broomhandles W i d o w cleaning squeegee handles Cherry picker lifts
952 Commercial and industrial applications of composites The considerable weight reduction com- Institute of Standards and Technology (NIST) pared with steel rolls means less strain and in Boulder, CO, and the Phillips Laboratory at fatigue and reduced injuries for the operators. Edwards AFB, CA, assist by conducting the For example, a steel roll that is slightly over low-temperature testing work. Competing 2489 mm (98 in) long, 152 mm (6 in) O.D., and technologies for SMES include pumped water 127 mm (5 in) I.D. would weigh approximately storage, batteries, and potentially, flywheels. Recent work has been focused on develop188 kg (239 lb). A fiberglass-reinforced composing and testing the prototype of a full-size ites roll of the same dimensions would weigh mockup of a SMES unit. E-glass fiber rein27 kg (61 lb), while a carbon fiber composites forced plastic curved, pultruded vinyl ester roll of this size weighs only 22 kg (49 lb). resin beams were developed to simultaneously provide essential electrical insulation 43.4.33 SMC - TOUGH, HIGH-PERFORMANCE (between conductor coils and between the Premix, Inc. (No. Kingsville, OH) and their conductor and the ground) and the conductor Quantum Composites division (Midland, MI), support structure. Pultruded FRP was chosen because it profocus on 'ultra-high strength and toughness' Lytex SMC that is offered with carbon fiber or vides reasonable structural performance and fiberglass reinforced epoxy or vinyl ester the required electrical insulation capabilities at resins. Present nonaerospace uses for this cryogenic temperatures for an affordable cost. unique SMC include a prosthetic foot, an auto- Test articles were fabricated and then the mobile brake component, an electrical mechanical, thermal, and electrical properties connector and more. Projected applications were measured over the temperature range of include automotive suspension and structural 4-295K. FRP beams, pultruded by Compositek chassis components, recreational-vehicle high- (of Brea, CA and now part of Kaiser impact parts and other applications Aerospace),met the goal of comparable perfordemanding toughness and excellent dynamic mance to currently used and more expensive materials such as G-10 laminate produced with fatigue resistance (including aerospace). fiberglass fabric and epoxy resin. The tensile failure strains for the materials tested using liq43.4.34 SUPERCONDUCTING MAGNETIC uid helium ranged from 1.2% at 295K to 1.9% ENERGY STORAGE SYSTEMS (SMES) at 4K. The pultrusion machine was built by Composites play a key role in the insulation Pultrex Ltd (Clacton-on-Sea, UK). Johnson and structural systems of the SMES Program Industries Composite Textile Reinforcements, at General Dynamics Space Magnetics, now a Inc. (of Phenix City, AL) supplied the stitched unit of Martin Marietta. Years of development fiberglass preform materials, which ranged may now offer dramatic potential for carbon from 9 to 18 plies in thickness. fiber and glass fiber composites in the future. SMES develops large radial loads when The purpose of SMES is to economically store charging. This requires strong coil support electrical energy from power production late straps, which are integrated by General at night and then to retrieve it during peak Atomics (San Diego) and which were filament demand periods during the day. wound by Lincoln Composites (Lincoln, NE). Heading the team is Bechtel National, Inc. The short straps look like a fan belt about as system integrator. General Atomics is 127 mm (5 in) wide and 762 mm (2.5 ft) long responsible for the coil supports and was the and are wound with S-2 glass fiber prepreg. designer and fabricator of the first SMES unit Larger straps, about 89-1200 mm (3.54 ft) to be operated successfully (1983) by an elec- long and 100 mm (4 in) wide, are wound using tric utility, Bonneville Power System. National carbon fiber prepreg.
Applications review 953 The SMES Engineering Test Model (SMESETM) consists of a two layer, radial, helically wound, 96 m (315 ft) diameter solenoid coil that has been designed to store a minimum of 20 MWh of electrical energy in its magnetic field. The coil stack requires 66 layers of FRP curved beams for the support structure or 825 24 m (80 ft) sections weighing 907 kg (2000 lb) each. A full-sized, 20 MWh ETM demonstration unit will require about 725 000 kg (1.6 million lb) of FRP composite. A full-scale test model of SMES will consist of a large, cylindrical coil 91 m (300 ft) in diameter sunk into a 9 m (30 ft) deep trench in the ground. Stored energy can be retrieved slowly or almost instantaneously, and the technology is considered a promising method of storing electrical energy for peak load demands. Bechtel estimates that the potential SMES market in the USA alone could be as high as $3 billion/yr. The Electric Power Research Institute in Palo Alto (CA) and the SMES Utility Interest Group (SUIG, Dallas, TX) are promoting the commercial development of the SMES technology because it looks promising for electrical load leveling, increasing electrical system reliability, savings in fossil fuels, reduced air pollution, emergency backup power, virtually instantaneous response and the ability to store energy from intermittent sources such as solar and wind generation of electricity. EPRI estimates that by 2010, storage could usefully provide as much as 10% of the nation’s generating capacity. Using storage to save power generated at night when power plants normally run at very low rates some day could potentially save the cost of building many new generating plants at far higher cost than adding SMES storage capacity.
and is produced using 68 w/o AS4 carbon fiber filament wound in PEEK thermoplastic matrix. They have reportedly produced well over 45 360 kg (100 000 lb) of parts with this material ’under production conditions, to aerospace requirements’. Also offered are parts molded from their injection-molding grade. Polygon offers a line of ’Poly Lube’ composite bearing materials, hybrids of fiberglass and Teflon. BalSeal Engineering Company (Santa Ana, CA) markets carbon fiber and graphite-filled Teflon fluid seals and ’0’-rings. The primary benefits are dimensional stability and a low coefficient of friction under elevated temperature and pressure, with no ’extrusion’ of the resin matrix. Carbon fiber adds to the tribological value of the rings through prevention of adhesion of the Teflon to the bearing surface. Tiodize (Huntington Beach, CA) sells selflubricating composites based on carbon fiber in a high-temperature resin used for bushings, thrust bearings, rod-end bearings and ballbearing retainer rings. They emphasize the advantage of using these materials compared to aluminum in retaining specific modulus at temperatures to 316°C (600°F). 43.4.36 URANIUM ENRICHMENT CENTRIFUGES
Louisiana Energy Services is a joint effort of three USA electric utilities, Duke Power Co., Louisiana Power & Light Co. and Northern States Power Co. Along with the internationally known engineering firm of Fluor Daniel, Inc., Urenco will build and help operate a uranium fuel enrichment plant for these utilities. Urenco is a European organization with close to 15 years of experience operating commercial uranium enrichment plants in Germany, the UK and The Netherlands. Together, the 43.4.35 TRIBOLOGICAL APPLICATIONS - SEALS three utilities own and operate 11commercial AND ‘O-RINGS nuclear power plants. In the enrichment EGC Corporation (Houston, TX) is a custom plant, high-speed ultracentrifuges will be fabricator of molded components for oil field used to enrich the uranium and hopefully can and refinery applications. Their ‘wear-resistant supply nuclear plant fuel at a price lower than composite’ is called XC-2 tribological material is being projected by DOE, according to the
954 Commercial and industrial applications of composites
program planners. Carbon fiber/epoxy filament wound composite tubes are to be used for the high-speed rotors but most of the technology details remain proprietary to Urenco. At this writing, Louisiana Energy Services continues with the approval process for the required environmental impact report in support of their license application. They hope to start construction soon, and have the new plant in full operation in the late 1990s. Similar composite centrifuge rotors are believed to be standard design for European and Japanese uranium enrichment work. 43.4.37 VALVES
Dresser Valve & Controls Division, Nil-Cor Operations, (Alliance, Ohio), sells several types of composite valves to the chemical industry. Over 100 000 Nil-Cor advanced composite valves are currently in service at over 100 USA paper mills. Their current valve product line includes fiberglass and carbon fiber reinforced PPS, polysulfone and vinyl ester in several sizes and types. Fiberglass reinforcement and Dow Chemical's Derakane 470 vinyl ester resin (among others) are used in ball and butterfly valves for the chemical industry. 43.4.38 WIND TURBINE BLADES
TPI, Inc. located their composite wind turbine blade production in Portsmouth, RI. TPI makes fiberglass/vinyl ester blades regularly for large operators of wind energy farms in the USA. The size of the individual generating units has gradually increased from the 50 and 100 kW sizes with 8 m (27 ft) composite blades that were most common several years ago to the larger 300-kW units that are now in demand. Each of the three new size blades is 17 m (56 ft) long and uses a sizable quantity of material. Production rates have reached as high as 5000-6000 blades per year. Among the critical aspects of making large windmill blades is the need for proper balance. This is one reason that the TPI product has succeeded
while other blade manufacturers had to fold when the market started to collapse several years ago. Sandia National Laboratories (of Albuquerque, NM) developed advanced wind turbine technology using the vertical axis approach and signed a technology agreement with Flo-Wind Corp. (Pleasanton, CA). FloWind has used extruded aluminum blades but plans use of composite blades. Flo-Wind currently operates 'over 500' such turbines with aluminum blades currently in operation on two California windfarms, likely candidates for retrofit with composite blades. Glasforms (San Jose, CA ) is one of their prime suppliers for composite parts. Competition continues to develop overseas. Japan's Tomen Corp. has joined with American and British firms to build several large wind power farms in Wales and Cornwall, UK, operated by a company called British Wind Farms, based in London. Included in their plans was the construction of 300 windmills with a total generating capacity of 77 000 kW. Many European countries are making sizable commitments to increased use of windpower, including Denmark, Germany, Holland and Spain. The National Renewable Energy Laboratory, a unit of the Department of Energy (formerly called the Solar Energy Research Institute), is in Golden, CO. They are interested in all types of alternate energy systems, including battery-operated vehicles, engines that can use different types of fuels, including hybrids and biofuels. NREL sponsored an effort called the 'Advanced Wind Turbine Next Generation Preliminary Design Project', with the goal of bringing wind power generation costs down from the current 7-84 to below the P 5 4 per kwh with coal generation. Wind generators must be made of materials likely to last at least 30 years, so the additional cost of using carbon or aramid fibers to reduce rotor blade weight and/or improve performance might be amortized over a long period.
Market data 955 Today there are over 15 000 windmills in California alone, with a total electrical generating capacity of 1300 megawatts. To date, they have generated a total of over two billion kwh, replacing 3.5 million barrels of oil. However, this is only 1.7% of the total power requirement of the state of California during the same period. There are more windmills overseas where about 70% of the total has been installed.
at twice the growth rate of the Gross Domestic Product (GDP). Sectors that fall mainly into the commercial and industrial applications arena, the focus of this chapter, are noted next. The corrosionresistant application sector for chemical storage, process tanks, pipe systems, cooling tower components, wastewater treatment, pressure vessels, and pollution control equipment grew 5.9% in 1993. This is because of their ties to capital spending and substitution 43.5 MARKET DATA for other materials, and another 3.7% gain is A frame of reference on market size is useful in likely in 1994.Electrical and electronic applicaassessing composites growth prospects. The tion usage grew at 5.7% in 1993 (plus 5.5% Composites Institute of the Society of the more in 1994). Power poles and fiber optic Plastics Industry issued a news release cable splice boxes are examples of growing February 8,1994 stating that USA shipments of areas here. The appliance, business equipfiber reinforced plastics (FRP) reached an esti- ment, and consumer product markets gained mated 1.23x lo9kg (2.73billion lb) in 1993. This only 3.0% (plus 3.7% more in 1994) and 2.2% was an increase of 6.9% over 1992 (thisis almost (with another 4.2% in 1994) respectively, entirely fiberglass-reinforced polyester, vinyl mainly from material substitution. Lastly, the ester, and epoxy material but includes 'other' category grew at 7.1% in medical advanced fibers composites). Projected 1994 equipment such as orthopaedic appliances growth to 1.32 x lo9 kg (2.90 billion lb) of com- and dental materials. The Composites Institute posites represents an increase of another 6.2%. (CI) represents over 415 firms in the composOverall growth in sales of composites continues ites industry and is the SPI's largest division. Table 43.10 USA composites shipments: 1984-1994 (in million Ib)
Markets
1984
1985
1986 1987 1988
1989 2990
1991
1992
1993
1994
projected (revised) 37 36 39 41 39 38.7 Aircraft/aerospace/ 29 32 military Appliance/business 123 133 137 141 150 151 153 135.2 equipment 420.0 Construction 430 445 456 506 495 470 468 Consumerproducts 143 142 149 167 169 158 165 148.7 Corrosion-resistant 310 295 291 329 349 335 350 355.0 equipment 231.1 Electrical/ 189 191 201 214 230 229 241 electronic 405 375 275.0 309 335 340 413 452 Marine 682.2 540 563 585 656 695 677 705 Transportation 80 76 79 73.8 83 75 Other 80 82 Total 2153 2218 2279 2537 2659 2542 2575 2360
32.3
25.4
24.6
143.2
147.5
153.0
483.0 162.2 332.3
530.0 165.7 352.0
575.0 171.2 365.0
260.0
274.9
290.0
304.4 750.0 83.4 2551
Includes reinforced thermoset and thermoplastic resin composites, reinforcements, and fillers. Source: SPI Composites Institute, 1994 Semi-annual statistical report, February 8,1994
319.3 332.8 822.1 890.0 89.3 93.9 2726 2895.5
956 Commercial and industrial applications of composites 43.6 PREDICTIONS
Much of the long-term growth for composites is likely to come from this miscellaneous sector of the business. The most likely segments to experience large unit growth rates are CNG tanks, flywheels, industrial rolls, distribution and lighting poles, civil structures, consumer and medical products and some other oil and gas or transportation components. Composites have a low market share in many industries, and commercial and industrial applications should continue to experience well above average growth compared with other composites, other material systems, and the total economic growth rate on a worldwide basis. Such growth over the long term is believed likely to exceed the 5-8% growth rate expected for 1994 over 1993 shipments for these composites markets. 43.7 CONCLUSIONS
Because these sectors of business and technology are highly fragmented many smaller producers are likely to participate. The scale of production is often not well suited to large producers. Also, the cost for large producers to pursue such fragmented businesses, pieces of which may not be very large, could be higher per unit or unit weight than for their huge composites commodity businesses such as automotive, marine and others. Much of the innovation in material forms and process technology also resides in this group of markets.
Thus, one should look to these markets not only for growth but also for technology and new ideas. REFERENCES
Most of the information provided in this chapter was generated through attendance at trade shows, interviews with industry participants, review of press releases, other extensive reading of publications, and the experience of the author over 30 years. Therefore, since most paragraphs would have required almost a reference per sentence, we elected to list below the best generic and reliable USA-based sources available to research these applications and markets or to track their future activities, participants and prospects. Composites Institute of the Society of the Plastics Industry, New York, NY. Composites News: Infrastructure and Advanced Material newsletters. Loud, Steve, (ed.) Solana Beach, CA: Composites Worldwide, Inc. Composites Industry Monthly and ACM Monthly newsletters. Burg, Martin, (ed.) San Diego, CA: Composite Market Surveys. Composites Technology and High-Performance Composites magazine. Hazen, Judith Ray, (publisher), Denver, C O Ray Publishing, Inc. Performance Materials and Fleets t? Fuels newsletters. Piellisch, Richard, (ed.) San Francisco, CA: August Pacific Press. plasticsBRlEF Reinforced Plastics Newsletter. Best, James R., (ed.) Toledo, OH: Market Search, Inc.
COMPOSITE BIOMATERIALS
44
Shalaby W. Shalaby and Robert A. Latour
44.1 COMPOSITES AS A FAMILY OF BIOMATERIALS
phosphate-based fiber showed large increases in both strength and stiffness (Casper et al., 1985; Lin, 1986). However, these families of composites lose a substantial fraction of their strength while retaining greater proportion of their stiffness after short term exposure to an aqueous physiological environment. Attempts to improve the strength retention have included coating the composite material to retard the transport of fluids to the polymer/fiber interface (Kelley et al., 1988; Andriano, Daniels and Heller, 1991) and surface modification of the phosphate fibers with a siloxane film barrier (Andriano, Daniels, and Heller, 1992).Growing interest in the inorganic phosphate-based fillers led Andriano and coworkers to compare the biocompatibility of several phosphate, fiber-reinforced polymers in a preliminary study (Andriano et al., 1993). Thus, phosphate fibers of calcium-sodium metaphosphate (CSM), sodium-calcium-aluminum-polyphosphate (NCAP) and potassium metaphosphate (PMS),with copolymers of E-caprolactone and lactide or polyorthoesters have been used as organic matrices. The CSM and NCAP fibers were found to be acutely nontoxic in cellular tissue and whole animal evaluation.
Although the use of high modulus metallic devices for internal bone fixation has been successful, there is a call to develop nonabsorbable, polymeric, composite substitutes having elastic modulus approaching or slightly higher than those of bones. This is to alleviate the problem of stress-shielding and consequent bone resorption as well as any concerns related to toxicity of metallic ions as corrosion products. (Gillett et al., 1984; Tonino and Folmer, 1987). Furthermore, to avoid a second surgery to remove a non-absorbable device and to allow for gradual load transfer to healing bones, completely absorbable composites were proposed and evaluated by many investigators (Casper et al., 1985; Daniels et al., 1990; Lin, 1986; Tormala et al., 1991; Vainionpaa, 1987). Among the key types of totally absorbable composites are those based on polyesters or polyorthoesters, reinforced with organic fibers (Tormala et al., 1991; Vainionpaa et al., 1987) or calcium or sodium/calcium polymetaphosphates (Casper et al., 1985, Andriano et d., 1993). A family of composites having organic polyester fibers as fillers (e.g. poly-L-lactide reinforced with polyglycolide fibers, displayed large increases in strength and modest increases in stiffness (Tormala et al., 1991; 44.2 COMPOSITE BIOMATERIALS: GENERAL Vainionpaa et al., 1987). On the other hand, polylactide reinforced with inorganic calcium Like the more traditional composites, those classified as composite biomaterials contain two or more distinct constituent materials or Handbook of Composites. Edited by S.T. Peters. Published phases on a microscopic or macroscopic size scale but not on the atomic level. Thus, fiberin 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
958 Composite biomaterials glass-reinforced polymeric composites and other reinforced resins are composites while metal alloys are not. Interest in the former systems was associated with the need to replace or augment biological tissues which are composites in their own rights. For instance bone, skin and blood vessels are typical natural composites. In both natural and man-made composites (or composite biomaterials) the strength and modulus of the matrix are dependent on the shape, stiffness and orientation of the reinforcing materials as well as their adhesion to the matrix. High modulus fibers uniaxially oriented in a low modulus matrix can produce stiff solid composites with maximum strength and modulus along the fiber direction as in carbon fiber-reinforced epoxy resins. Complex high density natural composites include cortical bones, dentin, cartilage as well as wood. Should air be a distinct phase of the composite, a foam is produced, as in synthetic sponges or lung tissues and cancellous bone (Park and Lakes, 1992). Early application of hard or dense synthetic composites as biomaterials includes the use of reinforced acrylics as dental fillings. The composite resins consist of a polymer matrix such as crosslinked methacrylate resin and a stiff, inorganic filler including barium glass or silica. The methacrylate resin can be based primarily on the glycidyl methacrylate derivations of bisphenol A. The choice of such components is consistent with the repair site, i.e. dentin, which in turn is a composite, made primarily of collagen and microscopic inorganic crystals of an apatite of calcium and phosphate. In general, the primary tissue of the teeth such as dentin is denoted as mineralized tissue where the primary function is load-bearing. A closely related tissue to dentin, both positionally and functionally is bone, which is a complex mineralized collagen. The composition and orientation of the bone components as loadbearing composites vary in different bones. Thus, in designing synthetic devices for repairing bone, one has to take into account the properties of the hard tissue in question. These
include strength, creep stiffness, and fatigue characteristics. For example, for anchoring a metallic hip prosthesis a high modulus crosslinked, filled methacylate 'grouting' or bone cement is usually sought. For cementless hip prostheses, carbon-fiber reinforced high performance thermoplastic polymers, such as poly(ether-ether ketone), PEEK, or aromatic poly sulfones (PS) are being explored. Cartilage is a low-load bearing, natural composite and its synthetic substitutes can be made of reinforced elastomers. Elastomeric tissues such as skin and blood vessels are made primarily of collagen and other biopolymers which impart compliance and elasticity. Synthetic substitutes of these tissues are far from being perfected. However, a key consideration is the use of an elastomeric polymer or design to provide inherent or engineering elasticity, respectively. Bioceramics and particularly hydroxyapatite (HA) have been used or proposed for use in many orthopedic and dental applications (Heimke, 1989, 1990). However, the mechanical properties of HA were not sufficient for demanding load bearing applications, such as fracture fixation or spinal fusion, and this led to its limited use in this area (Bostman et al., 1989). To address this issue, Knowles and Bonfield (1993) developed a glass-reinforced HA with enhanced mechanical properties. Utilizing glasses of the types xNa,O-(l-x) P,O, and xCaO-(l-x) P,O, (where x = 0.2, 0.3 and 0.5), a systematic study was conducted on the effect of increasing network-modifying oxides in the glasses on the mechanical properties of composites containing 2.5 and 5 wt Yo filler. The results indicate that HA reinforced with CaO-P,O, glass at about 2.5 wt Yo is of significant benefit to both the stabilization of HA and production of high flexural bend strength composites. This is illustrated in Fig. 44.1. for three types of CaO-P,O, glasses where that mole ratio CaO/P,O, was 20/80, 30/70 and 50/50 for composites designated as C,P, C,P and C,P, respectively.
Orthopedic implant applications 959 nents for joint arthroplasty. Each of these applications has a unique set of material and . 2.5YOC3P mechanical requirements which are addressed -x2.5YG5P 120 in this section. 100 FRP composite materials can be consida P ered to be composed of at least three, and z 80 u) possibly four, distinct components: the fiber ftl 60 which reinforces the matrix; the polymer 40 matrix which provides three-dimensional 20 support to the fiber; and the fiber/matrix interface which serves as the agent of load o i 1150 1203 1250 1303 1350 1403 transfer between the fiber and polymer Flrlng Temperature (C) matrix. In certain formulations, a fourth component, the fiber/ matrix interphase, must Fig. 44.1 Effect of glass composition on flexural bend strength for three different glasses at 2.5% also be considered. The interphase represents wt YO additions (from Knowles and Bonefield, the volume of matrix immediately adjacent to each fiber which is influenced by the fiber, 1993). causing the matrix in this area to have morphological differences (i.e. crystallinity) 44.3 ORTHOPEDIC IMPLANT APPLICATIONS and/or compositional differences (i.e. mixThe incentive for use of fiber reinforced poly- ture of matrix and sizing agent from fiber) mer (FRP) composite materials in most compared to the bulk matrix material. The nonmedical applications is provided by their site-specific biologic environment can potensuperior specific mechanical properties (i.e. tially influence each of these components of strength/weight, stiffness/weight ratios). the composite differently. Fortunately, experiHowever, this offers very little benefit for ence has shown that the in vivo environment orthopedic implant applications. Implants are can often be suitably represented by simple typically sufficiently small such that weight is physiologic saline solutions at 37°C (99°F) not an important design requirement. and pH = 7.4 for in vitro biomaterials perforHowever, there are other very important med- mance investigations. However, it must be ically related incentives for the development recognized that exceptions to this are not of these materials for the treatment of muscu- uncommon, and site-specific in vivo testing is essential prior to clinical evaluation of new loskeletal problems. While there are numerous potential ortho- composite material formulations; not only for pedic applications for FPR composite implant biocompatibility assessment, but also to materials, most work has been concentrated in ensure materials response to the biologic four specific areas. FRP composite materials environment has been properly understood. are being investigated for the design of FRP composites are uniquely different comfemoral components for total hip arthroplasty pared to metals in that they are permeable to and for pins, plates, screws, and nails for frac- moisture and salt ions. Therefore, if environture fixation as alternatives to metal alloys. mental durability is to be properly Fiber reinforcement is also being investigated investigated, test samples ideally should be as a means of improving the fatigue resistance fully saturated in their test environment prior of polymethylmethacrylate (PMMA) bone to testing. The American Society for Testing cement and the wear, creep, and fatigue resis- and Materials (ASTM) F04 Subcommittee on tance of ultrahigh molecular weight Composite Materials is currently developing polyethylene (UHMWPE)articulation compo- standards for environmental conditioning 160 140
-
960 Composite biomaterials prior to materials performance evaluations to address this issue. The biologic response to an implant material is as important as the material response to the environment. Biologic response to an implant material has been found to be strongly dependent upon whether the material in question can be internalized (phagocytosed)by the cells of the body which regulate the body’s foreign body response. Thus, a material which is very well tolerated in bulk form when implanted in the body (i.e. too large to be ingested by cells) may elicit a strong inflammatory response when in particulate form if the material particles are sufficiently small (i.e. < 10 mm (0.4 in)) to be phagocytosed by cells which mediate inflammatory response (Goldring, Clark and Wright, 1993; Black, 1992). Thus, in the development of composite biomaterials, it is important to not only assess the biocompatibility of the device in question, but to also assess biologic response to wear debris and degradation products of the fiber and polymer matrix materials. 44.3.1 FEMORAL COMPONENTS FOR TOTAL HIP ARTHROPLASTY
Femoral components are permanent implants ideally intended to last the entire life time of the patient. Current day metallic devices for this application have reported life expectancies of 90% survival out to 10-15 years in the elderly patient population, lower survival rates in younger patients (Callaghan, 1990). While this is considered good performance for an orthopaedic implant, certainly 90% survival at 30 years would be much preferred. Metallic components have the draw-back of being so stiff that much of the applied joint load is bypassed around the normally highly stressed proximal-medial region of the femur (calcar). This condition has been documented to lead to stress-shielding induced bone resorption (Sumner et al., 1992). In severe cases, this may not only contribute to component failure through loosening, but may also
severely complicate revision surgery (Engh and Bobyn, 1988). Several theoretical and experimental investigations have indicated that more compliant femoral components result in higher stress levels in the calcar area of the femur, thereby reducing stress shielding and maintaining greater levels of bone stock mass and quality (Bobyn et aI., 1992; Huiskes, 1992). While this is true, a widely reported misconception concerning this point is that an optimal condition would be reached if a femoral component could be designed to match the longitudinal elastic modulus or stiffness of cortical bone. This idea is completely erroneous and represents a failure to understand the mechanics of femoral component/femur load transfer. As long as a femoral component relies on intramedullary fixation, non-physiologic load transfer will occur, leading to some degree of stress shielding in the proximal femur. However, as a general concept, the degree of stress shielding should decrease proportionately to stem bending compliance. Three mechanical factors must be considered for the successful design of FRP composite femoral components. These are: (1) increased stem compliance to reduce calcar stress shielding; (2) fixation to bone and maintenance of acceptably low stem/bone interfacial stresses to prevent loosening; and (3) maintenance of acceptable stress levels within the femoral component to prevent fatigue failure. These factors are not independent; both interfacial stress level and component fatigue strength will be influenced by implant compliance. An optimal design may exist which can satisfy all three mechanical requirements and provide a superior hip joint replacement over current day metallic devices. Because of low bending stiffness, FRP composite stems cannot be properly evaluated by test methods developed for stiff metallic stems. The ASTM is currently developing a standard practice for fatigue testing of compliant FRP composite stems.
Orthopedic implant applications 961 Two FRP material systems are primarily being considered for femoral component development: carbon fiber-reinforced polysulfone (CF/PSF), and carbon fiber-reinforced polyether etherketones (CF/PEEK) (Davidson, 1987; Skinner, 1988). Several studies have addressed the effect of the biologic environment upon the mechanical material properties of these two composite materials with mixed results. CF/PEEK has been demonstrated to be very durable in physiologic saline environments when PEEK was APC2 grade (ICI, Tempe, AZ) (Strait et al., 1991; D’Ariano et al., 1994;Zhang et al., 1994),while 380 grade PEEK (IC1 Films Inc., Wilmington, DE) has been found to be sensitive to hydrolytic degradation of the fiber/matrix interface (Meyer and Latour, 1993). CF/PSF strength has been reported in separate studies to be both very stable (Overland et al., 1993) and significantly degraded by exposure to physiologic saline and exudate environments (Strait et al., 1991; Latour and Black, 1992,1993).Biocompatibility studies of both CF/PSF and CF/PEEK composite materials have suggested that, in bulk form, these materials should provide acceptable biocompatibility for use in femoral component applications (Wen et al., 1990). However, the generation of wear debris from implant/bone abrasion, and subsequent potential inflammatory reactions, is an important concern. 44.3.2 FRACTURE FIXATION DEVICES
FRP composite material devices are being developed for the replacement of metallic plates, screws, pins, and nails for fracture fixation. In contrast to joint replacement, these applications require only temporary implants. The ideal fracture fixation device would be sufficiently rigid initially to provide adequate fracture stabilization, then gradually decrease in stiffness over time to transfer stress to bone after fracture union to avoid stress shielding, and then eventually disappear to eliminate the need for retrieval surgery.
Two classes of FRP composite fracture fixation devices have been primarily considered: nonabsorbable and fully absorbable. Compliant nonabsorbable plates and nails have been investigated as a means of avoiding stress shielding associated with the use of stiff metallic components (Woo et al., 1976). Unfortunately, clinical results have indicated that, in high load bearing applications such as the tibia, compliant fracture fixation plates allow excessive motion at the fracture site causing unacceptable levels of pain upon weight bearing (Tayton et al., 1982; Tayton and Bradley 1983). This presents a ’Catch-22’ situation in which compliant bone plates may only be able to be utilized in very low load bearing applications, however, without high load bearing, stress shielding is not a serious concern and compliant plates are then no longer needed. Whde the advantages of nonabsorbable composite fracture fixation devices are therefore questionable, fully absorbable composite fracture fixation devices offer significant potential advantages over currently used metallic devices. These materials have the potential to be developed into components with sufficient initial stiffness and strength for load bearing fixation, which are then slowly degraded and absorbed by the body following healing such that implant retrieval is unnecessary. Two types of fully bioabsorbable composites have been investigated: (1) polymer fiber reinforced polymer, and (2) ceramic reinforced polymer. Bioabsorbable polymeric materials for both fiber and/or matrix which have been widely investigated are polyglycolic acid (PGA), polylactic acid (PLA), polydioxanone (PDS) (Tormala et al., 1991; Bostman et al., 1991), poly-e-caprolactone (PCL) (in’t Veld, 1993), and polyhydroxybutyrate (PHB) (Knowles et al., 1992). More recently, bioabsorbable polycarbonates and ’pseudo’polyamino acids have also been developed as well (Pulapura and Kohn, 1992). Types of ceramic fibers which have been investigated are phosphate glasses and calcium phosphate based ceramics (Andriano, Daniels and Heller,
962 Composite biomaterials 1992). Implant designs have utilized both continuous and discontinuous fiber reinforcement of the polymer. The main obstacle to widespread application in these materials is their rapid loss of strength and stiffness in vivo. This currently restricts their use to only low load bearing applications. This behavior has primarily been attributed to rapid hydrolysis of the fiber/matrix interface (Andriano, Daniels and Heller, 1992). The development of more durable fiber/matrix interfacial bonding in absorbable composites is essential if these materials are to be successfully developed for load bearing applications. The biocompatibility of bioabsorbable polymers and fibers being developed for orthopaedic applications is considered to be good, however, clinical use of self-reinforced PGA pins has demonstrated an 8% rate of aseptic sinus tract development in patients (Hofmann, 1992). While not compromising fracture union, this may require surgical intervention. Experience suggests this may occur when implant degradation product generation exceeds the local tissue clearance capability. The balance between degradation product release and tissue clearance ability raises a concern for the development of larger implants of these materials. This problem may potentially be overcome with the development of more slowly degrading implants.
devices typically occurs by loosening via combined fatigue fracture of the cement and the bone/cement or cement/implant interfaces. Fiber reinforcement is a potential means of improving the fatigue strength of bone cement. Several investigations have addressed this issue with the use of short fiber reinforcements of carbon (Pilliar et al., 1976), polyaramid (Wright and Trent, 1979), ultrahigh molecular weight polyethylene (Wagner and Cohn, 1989), titanium (Topoleski et al., 1992)) stainless steel (Fishbane and Pond, 1977), and PMMA (Buckley, 1991), to name a few. In general, these studies have demonstrated the expected result that short fiber reinforcement can increase both strength and toughness of bone cement. Problems with fiber wetting, fiber distribution, void content, and increased viscosity are cited as the major problems which have prevented clinical implementation. Bulk bone cement is actually a self-reinforced particulate filled composite material. The composite formulation with greatest potential may therefore be the replacement of the PMMA microspheres with similar quantities of PMMA fibers. This may enable mechanical properties to be improved while not causing significant increases in cement viscosity during cure. 44.3.4 ARTICULATION COMPONENTS
Ultrahigh molecular weight polyethylene (UHMWPE) is extensively utilized in total Polymethylmethacrylate (PMMA) is utilized joint replacement prostheses to provide a low extensively in orthopedic surgery as a friction surface for articulation against a method of prosthesis fixation in joint replace- matched metallic or ceramic component. Wear, ment. Bone cement is prepared in the creep, and fatigue resistance are the major operating room by the surgical team by mix- problems associated with the use of plain ing methylmethacrylate monomer with UHMWPE in these application, especially for polymethylmethacrylate microbeads. The knee joint prostheses (Connelly et al., 1984). mixture is first allowed to partially polymer- Fiber reinforcement offers a mechanism of ize, and then is placed in a prepared surgical potentially improving these properties. site within a bone cavity. PMMA is a brittle Reinforcement with carbon fibers was inipolymer and can be considered the weak link tially considered for this application, and even in joint replacement. The failure of cemented utilized clinically for both knee joint and hip 44.3.3 BONE CEMENT
Composites for soft tissues 963 joint replacement (Wright et al., 1988). This, however, proved to be a very poor choice of reinforcement. Wear resistance studies with this material provided mixed reports ranging from significantly decreased to significantly increased wear rates (McKellop et al., 1981), while fatigue resistance was found to be actually decreased by an order of magnitude compared to the unreinforced UHMWPE (Connelly et al., 1984). This behavior can be explained by the very brittle nature of carbon fiber and the very low interfacial bond strength between the fiber and matrix (Meyer and Latour, 1991). This combination leads to rapid crack initiation and propagation and third body wear during articulation via fiber fragment release. This incident in the history of implant design provides a clear example of the complexities of composite materials behavior, and demonstrates the fact that fiber reinforcement of a polymer does not necessarily improve mechanical performance. Although carbon fiber was not a good choice for the reinforcement of UHMWPE, as the old saying goes, ‘the baby should not be thrown out with the bath water’. Potential improvement of wear, creep, and fatigue resistance of UHMWPE may still be achieved using other types of reinforcement. In particular, a reinforcement is required which has properties of high strength, high strain to failure, high bond strength to UHMWPE, high wear resistance, and which is as biocompatible as UHMWPE in both bulk and particulate form. While UHMWPE is the most commonly used polymer for articulation, other composite systems have been investigated such as carbon fiber reinforced triacine resin and polymer fiber reinforced elastomeric composites (Harms, 1984; Sutphin et al., 1993). Despite the failure of CF/UHMWPE articulation components, the reinforcement of polymeric articulation surfaces is still an active area of research for the development of improved prostheses.
44.4 COMPOSITES FOR SOFT TISSUES
Composite biomaterials have been mostly associated with their use in conjunction with hard tissues. However, composites for repairing or replacing soft tissues can potentially become quite important due to the growing interest in vascular and skin grafts, as well as bioartifical organs. Simple devices such as sutures for soft tissue repair have been patented in composite forms. For instance, composite silk sutures with compliant copolyester matrix have been described as having lower tissue reactivity, and higher strength retention in the biologic environment as compared with wax-coated silk sutures (Shalaby, Stephenson and Schaap, 1984). Composite, woven vascular prosthesis with absorbable [10/90 poly(^-lactide-coglycolide)] and nonabsorbable (polyethylene terephthalate) segments, were made and used as model systems for studying the derivation of neointima in vascular grafts (Greisler et al., 1988). Composite artificial blood vessels were prepared by injecting water-soluble chitosan derivatives (e.g. hydroxypropylchitosan) and heparin into a microporous polytetrafluoroethylene tube followed by freeze-drying (Yamamura et al., 1992).What may be considered as a bioartificial vascular graft is the one prepared by endothelial cell seeding of a woven synthetic graft using filtration technique (Idezuki, 1993). Thus, canine venous endothelial cells which were seeded onto a low porosity vascular prosthesis were allowed to grow around the fibers in early periods. They then formed a monolayer on the internal surface of the tube at later periods in vitro. Properties of polyester fiber blends were evaluated as totally (Greisler, et al., 1988a) or partially (Yu and Chu, 1993) absorbable vascular grafts. Melt-blended absorbable polymers made of lactide and glycolide having the proper chemical composition to provide controlled miscibility in the liquid state were molded into components of surgical staples having the
964 Composite biomaterials desired in vivo strength retention profile (Smith et al., 1988;Jamiolkowski et al., 1989).In the solid state the molded articles exhibited a two-phase morphology. The texture of the dispersed phase m a y allow one to denote these systems a s microcomposites.
Casper, R.A., Kelley, B.S., Dunn, R.L., Potter, A.G. and Ellis, D.N. 1985. Fiber-reinforced absorbable composites for orthopedic surgery. Polymer Mater. Sci. Eng. 53: 497-501. Connelly, G.M., Rimnac, C.M., Wright, T.M., Hertszberg, R.W. and Manson, J.A. 1984. Fatigue crack propagation behavior of ultrahigh molecular weight polyethylene. J. Orthop. Res. 2: 119-125. REFERENCES Daniels, A.U., Chang, M.K.O. and Adriano, K.P. 1990, Mechanical properties of biodegradable Andriano, K.P., Daniels, A.U. and Heller, J. 1992. polymers and composites proposed for internal Biocompatibility and mechanical properties of a fixation of bone. J. Appl. Biomater. 1: 57-78. totally absorbable composite materials for orthopaedic fixation devices. J. Appl. Biomat. 3: DAriano, M.D., Latour, R.A. Jr., Kennedy, J.M., Schutte, H.D. Jr. and Friedman, R.J. 1994 Long 197-206. term shear strength durability of CF/PEEK Andriano, K.P., Daniels, A.U. and Heller, J. 1992. composite in physiologic saline. Trans. SOC. Mechanical Properties of Composites Reinforced Biomat., 17, 184. with Suvface-ModifiedAbsorbable Calcium-SodiumMethaphosphate Microfibers. Proc. 4th World Davidson, J.A.. 1987. The challenge and opportunity for composites in structural orthopaedic Biomat. Congr., Berlin, Germany. applications. J. Composites Techn. Res., 9(4): Andriano, K.P., Daniels, A.U. and Heller, J. 1991. 151-161. Biocompatability and mechanical properties of totally absorbable composite material for Engh, C.A. and Bobyn, J.D. 1988. The influence of stem size and extent of porous coatings on implant use. Trans. SOC.Biomater. 14: 10. femoral bone resorption after primary cementAndriano, K.P., Daniels, A.U., Smutz, W.P. and Wyatt, R.W.B. 1993. Preliminary biocompatibilless hip arthroplasty. Clin. Orthop. 231: 7-28. ity screening of several biodegradable Fishbane, B.M. and Pond Sr., R.B. 1977. Stainless of polysteel fiber reinforcement phosphate fiber reinforced polymer. J. Appl. Biomater. 4(1): 1-12. methymethacrylate. Clin. Orthop. Rel. Res. 128: 194-199. Black, Jonathan. 1992. Biological Performance of Materials. Fundamentals of Biocompatibility. New Gillett, N., Brown, S.A., Dubleton, J.H. and Pool, York: Marcel Dekker Inc. R.P. 1985. Biomaterials 6 113. Bobyn, J.D., Mortimer, E.S., Glassman, A.H., Engh, Goldring, S.R., Clark, C.R. and Wright, T.M. 1993. C.A., Miller, J.E. and Brooks, C.E. 1992. The problem in total joint arthroplasty: Aseptic Producing and avoiding stress shielding. loosening. J. Bone Joint Surg. 75A: 799-801. Laboratory and clinical observations of nonce- Greisler, HI!,Dennis, J. W., Endean, E.D., Ellinger, mented total hip arthroplasty. Clin. Orthop. 274: J., Buttle, K.F. and Kim, D.U. 1988 (a). 79-96. Derivation of neointima in vascular grafts: Bostman, O.M. 1991. Current concepts review. Circulation 78 (3 pt. 2) 1-6-12. Absorbable implant for the fixation of fractures. Greisler, H.P., Endean, E.D., Klosak, J.J., Ellinger, J., J. Bone Joint Surg. 73-A: 148-153. Dennis, J.W., Buttle, K. and Kim, D.U., 1988 (b) Bostman, O., Hirvensalo, E., Vainionpaa, H., Polyglactin 910/polydioxanone bicomponent Makela, A., Vihtonen, K., Tormala, P. and totally absorbable vascular prosthes. J. Vasc. Rokkanen, P. 1989.Ankle fractures treated using Surg. 7: 697-705. biodegradable internal fixation. Clin. Orthop. Harms J., Mittelmeier H., and Mausle, E. 1984. 238: 195-203. Results of animal studies on the use of carbon Buckley C.A., Lautenschlager E.P. and Gilbert J.L. fiber-reinforced plastic protheses. In The 1991. High strength PMMA fibers for use in a Cementless Fixation of Hip Endoprostheses. (ed. E. self-reinforced acrylic cement: Fiber tensile Morscher), pp. 249-251. New York: Springerproperties and Composite toughness. Trans. SOC. VerIag. Biomat. 14: 45. Heimke, G. 1989. (Ed.) Bioceramics Vol. 1, Ishiyaku Callaghan, J.J. 1990. Total hip arthroplasty. Clinical Euro-America, Japan. perspective. C h . Orthop. 276: 3340. Heimke, G. 1990. (Ed.) Bioceramics, Vol. 2, Cologne,
References 965 Germany, Deutsche Keramische Geselleschaft. Hofmann, G.O. 1992. Biodegradable implants in orthopaedic surgery - A review on the state-ofthe-art. Clinical Materials. 10:75-80. Huiskes, R. 1992. The relationship between stress shielding and bone resorption around total hip stems and the effects of flexible materials. Clin. Ortkop. 274: 124-134. Idezuki, Y., Shindo, S. Shirakawa, M. and Egami, J. 1993. N m Functional Materials, Vol. B. (Tsuruta, T. ed.), pp. 333, Elsevier, Amsterdam. in't Veld, P.J.A., Dijkstra, P.J., and Feijen, J. 1993. In vitro degradation of polyesteramides containing poly-caprolactone blocks. Clinical Materials. 13: 143-147. Jamiolkowski, D.D., Gaterud, M.T., Newman, Jr., H.D. and Shalaby, S.W. 1989. Surgical Fastener Made from Glycolide-Rick Polymer Blends. US Patent (to Ethicon Inc.) 4 889 119 Kelley, B.S., Dunn, R.L., Jackson, T.E., Potter, A.G. and Ellis, D.N. 1988. Assessment of Strength Loss in Biodegradable Composite. Proc. 3rd World Biomater. Congr., Kyoto, Japan, 471. Knowles, J.C. and Bonfield. 1993. Development of a glass reinforced hydroxyapatite with enhanced mechanical properties. The effect of glass composition on mechanical properties and its relationship to phase changes. 1. Biomed. Mater. Res. 27(12): 1591-1598. Knowles, J.C., Hastings, G.W., Ohta, H., Niwa, S. and Boeree, N. 1992. Development of a degradable composite for orthopaedic use: in vivo biomechanical and histological evaluation of two bioactive degradable composites based on the polyhydroxybutyrate polymer. Biornaterials. 13: 491496. Latour, R.A. Jr. and Black, J. 1993. Development of FRP composite structural biomaterials: Fatigue strength of the fiber/matrix interfacial bond in simulated in vivo environments. 1. Biomed. Mater. Res. 2 7 1281-1291. Lin, T.C. 1986. Totally absorbable fiber-reinforced composite for internal fixation devices. Trans. SOC.Biomater. 9: 166. McKellop, H., Clarke, I., Markolf, K. and Amstutz H. 1981. Friction and wear properties of polymer, metal, and ceramic prosthetic joint materials evaluated on a multichannel screening device. 1. Biomed. Mater. Res. 15: 619-653. Meyer, M.R. and Latour, R.A. Jr. 1991. Fiber reinforcement of ultrahigh molecular weight polyethylene. Trans. SOC.Biomater. 14: 285. Meyer, M.R. and Latour, R.A. Jr. 1993. The long-
term durability of interfacial bonding in carbon fiber/polytheretherketone and carbon fiber/polysulfone composites following exposure to simulated physiologic saline. Trans. SOC. Biomater. 16: 15. Overland, M.K., Clayden, N.J., Everall, N.J., Koeneman, J.B. and Magee, F.P. 1994. Effect of long-term in-vivo/in-vitro environmental exposure on the shear strength of polysulfone/carbon fiber composites. Trans. SOC. Biomater, 17, 159. Park, J.B. and Lakes, F.S. 1992. Biomaterials - An Introduction. 2nd ed., New York, Plenum Press. Pilliar, R.M., Blackwell, R., Macnab, I. and Cameron, H.U. 1976. Carbon fiber-reinforced bone cement in orthopaedic surgery. I. Biomed. Mater. Res. 10: 893-906. Pulapura, Sand Kohn, J. 1992. Trends in the development of bioresorbable polymers for medical applications. I. Biornaterials Applications. 6: 216-250. Shalaby, S.W., Stephenson, J. and Schaap, L. 1984. Composite Sutures of Silk and Hydrophobic Thermoplastic Elastomers. US Patent (to Ethicon Inc.) 4 461 298. Skinner, H.B. 1988. Composite technology for total hip arthroplasty. Clin. Ortkop. 235: 224-236. Smith, C., Gaterud, M., Jamiolkowski, D.D., Shalaby, S.W. and Newman Jr., H.D. 1988. Higk Glycolide Blends for Absorbable Staples. US Patent (to Ethicon Inc.) 4 741 337. Strait L.H., Jamison R.D., and Gavens, A. 1991. Effect of environment on the flexural and compressive strength of carbon/polysulfone and carbon/polyetheretherketone composites. Trans. SOC.Biomater., 14: 286. Sumner, D.R., Turner, T.M., Urban, R.M. and Galante, J.O. 1992. Experimental studies of bone remodeling in total hip arthroplasty. Clin. Orthop. 276: 83-90. Sutphin, C.M., LaBerge, M., Drews, M.J., Lickfield, G.C. and Black, J. 1993. Design of an elastomeric composite for orthopaedic applications. Trans. SOC.Biomater. 16:147. Tayton, K. and Bradley, J. 1983. How stiff should semi-rigid fixation of the human tibia be? A clue to the answer. J. Bone Joint Surg. 64-B: 105-111; 65-B: 312-315. Tayton, K., Johnson-Nurse, C., McKibbin, B., Bradley, J. and Hastings, G., 1984. The use of semi-rigid carbon-fiber-reinforced plastic plates for fixation of human fractures. Results of preliminary trials. J. Bone Joint Surg. 65-8: 312-315.
966 Composite biomaterials Topoleski, L.D.T., Ducheyne, P. and Cuckler, J.M. 1992. The fracture toughness of titanium-fiberreinforced bone cement. J. Biomed. Muter. Res. 26: 1599-1617. Tonio, A.J. and Folmer, R.C.H. 1987. The clinical use of plastic plates for osteosynthesis in human fractures. Clin. Muter. 2: 275-279. Tormala, P., Vasenius, J., Laiho, J., Pohjonen, T. and Rokkanen, P. 1991. Ultra-high-strength absorbable self-reinforced polyglycolide (SRPGA) composite rods for internal fixation of bone fractures: In vitro and in vivo study. J. Biomed. Muter. Res. 25: 1-22. Vainionpaa, S., Kilpikar, J., Laiho, J. Helevitro, P., Rokkanen, P. and Tormala, P. 1987. Strength and strength retention in-vivo of absorbable, selfreinforced polyglycolide (PGA). Biomuteriuls. 8: 46-47. Wagner, H.D. and Cohn, D. 1989. Use of high-performance polyethylene fibres as a reinforcing phase in poly(methylmethacry1ate) bone cement. Biomuteriuls. 10: 139-141. Wen, L.M., Merritt, K., Brown, S.A., Moet, A. and Steffee, A.D. 1990. In vitro biocompatibility of polyetheretherketone and polysulfone composites. J. Biomed. Mater. Res. 24: 207-215.
Woo, S. L.-Y., Akeson, W.H., Coutts, R.D., Rutherford, L., Doty, D., Jemmott, G.F. and Amiel, D. 1976. A comparison of cortical bone atrophy secondary to fixation with plates with large differences in bending stiffness. J. Bone Joint Surg. 58-A: 190-195. Wright, T.M. and Trent, P.S. 1979. Mechanical properties of aramid fibre-reinforced acrylic bone cement, J. Muter Sci. Letts. 14: 503-505. Wright, T.M., h a c , C.M., Faris, P.M. and Bansal, M.. 1988. Analysis of surface damage in retrieved carbon fiber-reinforced and plain polyethylene tibial components from posterior stabilized total knee replacements. J. Bone Joint Surg. 70-A: 1312-1319. Yamamura, K., Sakurai, T., Kizawa, H. and Harada, H. 1992. Jap. Pat. Appl. 92-58845; Chem. Abstr. 119(24):2565852. Yu, T.J. and Chu, C.C. 1993. Bicomponent vascular grafts consisting of synthetic absorbable fibers. J. Biomed. Muter. Res. 2 7 1329-1339. Zhang, G., Latour, R.A. Jr., Kennedy, J.M., Schutte, H.D. Jr. and Friedman, R.J. 1994. Long term compressive strength durability of carbon fiber reinforced PEEK composite in physiologic saline. Trans. SOC.Biomuteriuls, 17: 160.
SCIENTIFIC APPLICATIONS OF COMPOSITES 45 Vicki P. McConnell
45.1 INTRODUCTION
Science may be seen as a journey, encompassing both inner and outer space on the quest for greater understanding of the universe. This can take literal form, such as the journey toward ever farther, faster flight regimes of experimental aircraft like the X-30, or here on Earth aboard a high-speed train that levitates on magnetic rails. Or the journey can take theoretical form in scientific instruments that measure the smallest particles of matter on Earth and peer into the black holes‘ of space. From the human eye to ’eye in the sky’ space telescopes, science continually extends our vision of the universe. Science and composites have always intersected in the R&D laboratory with new materials and process discoveries, and in the fabrication of actual structures used in scientific applications. The material selected for the Hubble Space Telescope metering truss, for example, and the process used to build the components of that structure are directly related to its ability to accomplish a unique scientific mission. As one might expect, pure science is rarely accomplished in a vacuum; helpful data are drawn from commercial, industrial and military programs and may return to those venues if the scientific application involves technology transfer. Government and private sector sources may team up to make the scientific
Handbook of Composites. Edited by S.T. Peters. Published
in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
research possible. And while national technology agendas are largely determined by the presiding political administration, there is nothing temporary or partisan about the exponential growth of global competition. Government agencies in the USA have worked over the years to form a long-term national technology strategy that integrates materials science. These have included the National Critical Materials Council (established by Congress in 1984), the White House Office of Science and Technology (which coordinated President Bush’s National Materials Initiative and the $1.8 billion Advanced Materials and Processing Program), the National Center for Advanced Technologies (NCAT) formed in 1989 by the Aerospace Industries Association, the Department of Commerce’s Advanced Technology Program administered by the National Institute of Standards and Technology (NIST), and the Defense Advanced Research Projects Agency (DARPA) within the Department of Defense. Grants totalling nearly $500 million were awarded under President Clinton’s Technology Reinvestment Project (TRP) in this program’s first year, 1993. TRP channels defense department funds to projects that emphasize dual-use technology for both commercial and military applications, with the ultimate goal of enhancing cost and information sharing and ensuring that key USA industries stay on the technological leading edge (Lynch, 1993). The variety of compositerelated TRP projects have ranged from rapid densification of carbon-carbon components,
968 Scientific applications of composites infusion-molded infrastructure products, marine craft and dockside structure development to electric vehicle body and propulsion systems. 45.2 APPLICATIONS
For the scientific applications described in this chapter, carbon fiber-reinforced thermosets and thermoplastics offer a superior strength-to-weight ratio over metal and other materials, as well as extremely close fabrication tolerances, complex contours and load tailorability, part consolidation through integral design, and resistance to multiple degradation factors. In some cases, parts and structure simply could not be made to meet performance parameters without the use of advanced composite materials. 45.2.1 SCIENCE-MISSIONSPACECRAFT
For composites, the hard vacuum of space presents a rigorous environment which may include atomic radiation, thermal cycling, UV ionizing radiation and micrometeroid impacts. In answering the challenges of these synergistic effects on spacecraft, advanced composites provide crucial payload weight savings, nearzero CTE, and good dimensional stability when material formulation or design takes hygroscopy and outgassing into effect. Composites are commonly used in optical benches, precision mounts (including structural trusses), solar arrays and booms (Lubin and Dastin, 1982).
Spacecraft truss structure Carbon fiber/epoxy tubes joined into truss structures are commonly used for supporting instruments on science spacecraft, such as the propulsion subsystems and upper-stage assembly for the radar-mapping Venus probe Magellan. Some 60 tubes in the truss range in diameter from 25-100 mm ( 1 4 in) and in length from 180-1830 mm (7 in-6 ft). They
were designed for compression load capabilities from 860-6000 kg (1900-13 230 lb). The same thermoset prepreg used for Magellan parts worked well in the science and radioisotrope thermal-generator booms and low-gain antenna struts and mast on the Galileo Jupiter observer. Originally envisioned as one of four 'great observatories' in the sky, NASA's Hubble Space Telescope (HST)took its place in 1990 as the largest optical system in orbit - the size of a railroad boxcar weighing 11 567 kg (25 500 lb) (McConnell, 1989). Capable of studying wave lengths from far UV to far IR, the focal length of this 'eye in the sky' is 607 mm (189 ft). Pointing accuracy and stability are critical - the telescope may hold direction for as long as 24 h at a time depending on the object of focus (Fig. 45.1). Supporting the secondary mirror in precise alignment with the primary mirror is a carbon fiber /epoxy metering truss, engineered to maintain stability of the nominal 5080 mm (200 in) spacing between the mirrors within +15 x lo4 mm ( 6 0 microinches) in the presence of A1"C (e0"F) excursions at mean temperature of -73°C (-100°F). The truss was fabricated by Boeing Defense & Space Group (Huntsville, AL), along with a carbon fiber/epoxy focal-plane structure that supports five primary science instruments and three fine-guidance sensors. Additional composite HST parts include a support-systems module equipment shelf that holds three positioning gyros, and the fineguidance keel. Discrete internal ribs direct load paths out the main core members and prevent curvature in these parts. Shelf and keel were designed by Perkin-Elmer (Danbury, CT) and fabricated at Composite Optics Inc. (San Diego, CA). COI has developed a proprietary moisture-barrier plating technology to seal epoxies from the effects of moisture prior to launch. Both COI and Jet Propulsion Laboratory (Pasadena, CA) have built spacecraft test articles using a carbon fiber/cyanate ester (McConnell, 1992b) composite that shows less
Applications 969
Fig. 45.1 Composite structural truss (a) aboard Hubble Space Telescope (b)aligns primary and secondary mirrors. (Sources:NASA, Perkin Elmer.)
moisture absorption than epoxies and greater toughness than either epoxies or BMIs. Lastly, HST’s high-resolution spectrograph optical bench was manufactured by Hercules Aerospace Co. (Magna, UT) to specifications from Ball Aerospace Systems Group (Boulder, CO). Carbon fiber/epoxy offers CTE equal to quartz and maintains 2 m (6.5 ft) of optical surfaces to within 0.0254 mm (0.001 in). Launched in December 1995, the Solar and Heliospheric Observatory (SOHO) was designed to be pointed directly at the Sun to study the interactions between the Sun and Earth’s environment. The largest moving instrument aboard SOHO is the Ultraviolet
Coronograph Spectrometer (UVCS), designed as an occulted telescope to focus on solar winds (McConnell, 1993~).As with HST, optical mechanisms require accurate and repeatable positioning, so UVCS’s design centralized upon a three-segment carbon fiber/epoxy structure. Ball Aerospace Systems Group and COI built the seven-sided truss of flat panels that result in cylindrical shape. Some 80 separate stability requirements had to be met, along with strict mass requirements (the structure weighs 21.7 kg (48 lb) but supports 91.6 kg (202 lb) of instrumentation), and the ability to survive launch loads of 18G. High-modulus pitch-based carbon fiber /epoxy
970 Scientific applications of composites
prepreg met weight, mass and optical stability parameters with no outgassing or microcracking. Especially thin prepreg (0.063 mm (0.0025 in) cured ply thickness) also enhanced the thermal conductivity of the fiber, an added stability factor (Kilpatrick, 1992).
to the central cylinder, large curved and conical shell structure, mirror support sleeves, and solar array panels. Eastman Kodak built the 2643 long optical bench that forms the main structural element, with carbon fiber skins over honeycomb core. LDEF composite specimen testbed
AXAF orbiting observatory
Slated for launch from the Space Shuttle in August 1998, the Advanced X-Ray Astrophysics Facility spacecraft will bridge the gap between the two Great Observatories currently in orbit: Hubble Space Telescope and Compton Gamma Ray Observatory. TRW (Redondo Beach, CA) is prime contractor for AXAF and the spacecraft will contain four sets of unique cylindrical, grazing incidence mirrors. These are mounted concentrically in a nested array. Carbon fiber/epoxy is used in equipment compartments attached
Proof positive of the hostile effects of space startled cosmonauts on the Mir space station in 1992 when they found not a shred remained of the Soviet flag after a year in orbit on the station's exterior. Further proof comes in data retrieved from NASA's Long Duration Exposure Facility (or LDEF), a 12sided circular spacecraft launched into low Earth orbit (LEO) from the Challenger shuttle in April 1984 and retrieved in 1990 (Fig. 45.2). By 1993, postflight analysis of the 86 experiment trays containing more than 10 000 different material specimens has been carried
Fig. 45.2 End panel of LDEF spacecraft shows effects of space exposure on materials specimens. (Source:NASA.)
Applications 971 out by a team of international scientists. Research on LDEF's composites specimens has provided a benchmark for next-generation space materials, such as those that make their way onto the international space station (Stein, 1993). The six-year duration of LDEF's exposure to the LEO space environment provided the longest term flight data ever retrieved for analysis at that time. Post-retrieval symposiums have shared analytical conclusions, and NASA selected M/Vision software from PDA Engineering (Costa Mesa, CA) as one format for the sizeable LDEF Materials Database which contained research information on about 35% of the LDEF materials by the end of 1994. An important conclusion drawn from LDEF specimen analysis is that environmental effects are dependent upon the location of composite hardware on a spacecraft - leading or trailing edge, shielded or unshielded from degradation factors (only metal matrix composites with aluminum matrices showed no location-dependent degradation; magnesiummatrix composites oxidized on sample edges)
(Fig. 45.3). Atomic oxygen is identified as the most detrimental factor to polymer-matrix composites, especially on uncoated surfaces (George, 1992). Erosion has been observed to 0.127 mm (0.005in) (about one ply of laminate) on leading edge components made from carbon fiber with epoxy, polyimide and polysulfone matrices. Micrometeroid damage was not catastrophic on any LDEF specimens, but can compromise surfaces and lead to substrate erosion (Blair, 1992) (Fig. 45.4).Another conclusion drawn based on post-flight analysis was that predicting erosion rates and formulating material with enhanced A 0 resistance should key on fiber resistance as the dominant factor. A prediction model for A 0 erosion yield in polymer materials has been developed (Tennyson, 1993), based on the repeatable atomic composition of the polymer. An LDEF follow-on spacecraft, the European Retrievable Carrier (EURECA) was retrieved by the shuttle Endeavor in June, 1993 after 11 months in a 370 km (230 mile) orbit. EURECA's mission focused on materials and fluids in low microgravity and LEO exposure
Fig. 45.3 LDEF orientation in orbit. (Reprinted with permission of Bland A. Stein, NASA Langley Research Center.)
972 Scientific applications of composites 45.2.2 THE X-30 NATIONAL AEROSPACE PLANE (NASP)
This NASA/ Air Force single-stage-to-orbit technology demonstrator program begun in 1986 generated significant advanced research into aerothermodynamics, propulsion, and new enabling materials (Fig. 45.5). These accomplishments were due in great part to unprecedented teaming among competitive aerospace companies contracted for the project. NASP was designed as a piloted vehicle capable of horizontal takeoff and landing on conventional runways (with an airbreathing, hydrogen slush-fueled engine), hypersonic cruise (speed range to Mach 25), low Earth orbit, and operation at temperature extremes
Fig. 45.4 Two LDEF composite samples show micrometeroid damage in center. A 0 erosion affected only epoxy in E-glass/epoxy sample (a) but uniformly degraded matrix and fiber in carbon fiber/epoxy sample (b). (Source: Christopher Bland, Lockheed Missiles & Space Co.)
on solar array components. Shorter duration experiments (from 11 to 40 h of exposure) flown aboard the shuttle Atlantis in July 1992 focused on further analysis of atomic oxygen effects (particularly flux reaction and erosion rates) upon nearly 1500 material and coating samples. These experiments included NASA’s EOIM-3, and payloads sponsored by the University of Alabama and Case-Western Reserve University (Friebele, 1992).
Fig. 45.5 Materials science results in technology transfer from NASP demonstrator vehicle research (a) to a computer part that speeds up data access (b). (Source: NASk/Air Forch.)
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Applications 973
974 Scientifi'c applications of composites (cryogenic to 2760°C (5000°F)). R&D efforts resulted in innovative materials science among five composite categories (Mcconnell, 1990), with more than 2000 airframe and scramjet propulsion components, structure, and systems built (Table 45.1). Until October 1994, NASP represented the hypersonic flagship program of the USA. At that time, the initial program goals transitioned into the hypersonic system technology program (HYSTP), funded jointly by the US Air Force and NASA and which will focus on development of hypersonic scramjet propulsion. Yet resulting scientific R&D from the NASA program generated considerable technology transfer to a broad range of commercial applications, including oilfield pipe, auto engines and hip joint implants. One of the titanium MMCs developed did not exist before the NASP program and is 100 times more resistant to corrosion than standard aircraft titanium. An aluminum/beryllium alloy tested in NASP heat exchanger prototypes has since found application in computer actuator arms, providing a 20% increase in data access speed. Big Three automakers have studied NASP-developed titanium MMCs and alloys to cut weight by 50% in valves, piston rings, cam shafts, con-
Fig. 45.6 Composite aeroshell and nose cone enabled DC-X single-stage-to-orbit prototype launch vehicle to fly nose-first takeoff and tail-sitter landing. (Source: McDonnell Douglas Aerospace.)
necting rods and other engine parts. Thingauge titanium aluminide foil has been tested in heart valve assemblies and pacemaker cases. Adaptation of computational fluid dynamics (CFD) - the study of high-speed laminar air flow developed to maximize NASP's body configuration - has proven useful in analyzing aircraft and automotive aerodynamics, as well as blood flow in artificial hearts. Dr. Steve Charles of the Center for Retina Vitreous Surgery (Memphis, TN) began work with the Rockwell Science Center (Thousand Oaks, CA) in February 1992 to apply NASP-developed CFD to modeling fluid flow inside the eye. CFD will help analyze the role played by rapid eye motion (up to 600"/s) in retinal detachment, and possibly enhance the precision of surgical procedures. Dr. Charles performs some 15 000 vitreous retinal surgeries annually and ultimately expects to use CFD to design a new suite of surgical tools, 'applying hard science rather than guesswork to the entire tool set'. 45.2.3 DELTA CLIPPER EXPEFUMENTAL LAUNCH VEHICLE (DC-X SSTO)
Under a $60 million contract from the Ballistic Missile Defense Organization, McDonnell
Applications 975 Douglas Aerospace (Huntington Beach, CA) developed its Delta Clipper experimental prototype (DC-X) to prove the practicality, reliability, operatability and cost efficiency of a reusable single-stage-to-orbit (SSTO) flight vehicle. Comparative figures suggest that reusable rockets like a full-scale DC-X could reduce commercial/military payload delivery costs to under $500/lb - an attractive alternative to the $10 OOO/lb for the Space Shuttle (1993 figures). The one-third scale DC-X prototype had completed five suborbital flight tests before being grounded by an on-board explosion that damaged the composite aeroshell. Despite this accident, DC-X landed safely and the shell was repairable. Data collected in the first flight test series formed the advanced technology backbone for an upgraded DC-XA vehicle, developed jointly by McDonnell Douglas (now the Boeing Co.) and NASA (with some $43 million in NASA funding) in 1995. The DCX-A was renamed the Clipper Graham in 1996, and completed several milestone flights that year: altitude of 10 300 ft with 550 ft lateral movement, and a second flight 26 hours after the first to demonstrate rapid turnaround. Rapid prototyping played an essential role in the cost effectiveness of building the DC-X prototype, including construction of composite aeroshell, nose cone, and base heat shield components (McConnell, 1993~). The aeroshell and nose cones were fabricated by Scaled Composites (Mojave, CA); five aerodynamic composite flaps on the carbon fiber/epoxy aeroshell provided the ability to ascend nose first and return to Earth for a 'tail sitter' landing. Composites accommodated the program's ambitious scheduling with rapid aeroshell construction (nine months) and by allowing continuing modifications without huge cost overruns. Lay-up of woven fabric composite skins over foam core resulted in aeroshell wall thickness range of 12.7-38.1 mm (0.5-1.5 in). Besides enduring external temperatures to 149°C (300°F) and internal temperatures to
-73°C (-100"F), the aeroshell incorporated built-in longitudinal longerons to handle engine/truss, parachute deployment and landing gear loads. Glass fiber/epoxy was used in the 762 mm (2.5 ft) tall nose cone since it required signal transparency for antennae. McDonnell Douglas decided that carbon fiber/epoxy skins over syntactic foam in the DC-X base heat shield provided the best material system for meeting weight, price, and scheduling demands. The shield (about 396 mm (13 ft) on a side) featured a unique 'super circle' sectional design, joined with titanium fasteners. Both the prime and subcontractor worked to the philosophy of building real structures to collect and prove data before moving on to more expensive program goals, something of a first in such R&D efforts. The durability of composite aeroshell materials and overall system design was proven in the June 1994 explosion; according to the program director, the 1220by 4570 mm (4 by 15 ft) vertical tear would probably have demolished other launch vehicles. McDonnell Douglas developed and fabricated a robust carbon fiber/epoxy liquid hydrogen tank for DCX-A in 1996,8 ft in diameter and 16 ft tall, but 33% lighter than previous aluminium alloy tanks. 45.2.4 GROUND-BASED SCIENTIFIC INSTRUMENTS
Wind tunnel blade While CFD places air-flow analysis within the domain of computer mathematics, there is nothing quite like proving an aerodynamic concept in a wind tunnel. At NASA's Ames Research Center (Moffett Field, CA), differentsized wind tunnels can accommodate small models of experimental structures and shapes as well as full-scale vehicles (such as tractor trailers and an F/A-18 aircraft). Compressors generate high-speed air flow with rotating blades, usually constructed of aluminum or laminated wood. In the Ames tunnel with a
976 Scientific applications of composites 3352 mm (11ft) test section, composite design sandwich structure of syntactic foam core and prototyping work supported by Dr. Clem covered with hybrid glass and carbon Hiel, a former NASA senior design engineer, fiber/epoxy skins laid up at 0/+30/-30 oriand his associates has examined the potential entation. Though syntactic foam can typically of using composite blades in the compressor, be four to eight times higher in weight than which is capable of speeds to 711 rev/min. traditional foams, mechanical properties are These blades would offer the advantages of several orders of magnitude higher. This is improved damage tolerance, longer fatigue due to the energy-absorbing capacity of glass life, higher damping and improved safety in microspheres inside the foam core, which also localize damage when they are crushed. the event of catastrophic failure. Design of a prototype composite blade in In addition, local regions of skin failure are 1995 (Fig. 45.7) (which is 1067 mm (42 in) clearly visible to the unaided eye by an exterlong, 406 mm (16 in) wide at the root, taper- nal imprint. Carbon fiber skins served as the ing to 241 mm (9.5 in) at the tip) featured a structural backbone by providing high specific strength and stiffness, while the foam core supported skins against impact. Glass fiber skins acted as sacrificial protective coating and visual enhancement of impact damage during residual strength assessment. Dr. Hiel reports that the prototype composite blade tested to seven times the expected combination of centrifugal and air loads, and was clearly capable of carrying higher loads when the test fixture failed. Manufacturing the blade through resin transfer molding (RTM) could result in a blade that would compete with aluminum in acquisition costs while reducing weight significantly (Hiel, 1993). Particle accelerators
To study the exotic mixture of protons, neutrons, nuclei and the smallest particles of matter - quarks - inside the atom, physicists have turned to complex scientific instruments that generate enough voltage and acceleration to create high-energy particle collision. Analysis of the collisions in this 'quark soup' could reveal mysteries among the building blocks of matter. At Newport News, VA, the Southeastern University Research Association oversees operation of the Continuous Electron Beam Accelerator Facility (CEBAF). Here a continuFig. 45.7 Wind tunnel blade made of carbon fiber/epoxy skins over syntactic foam core (left) ous electron beam with up to 4 billion eV of weighs 50% less than aluminum counterpart. energy is steered by superconducting magnets through up to four orbits of an underground (Source:NASA Ames Research Center.)
978 Scientific applications of composites
ground tunnel in Ellis County, TX. The SSC design was based on a dual ring of 10 000 electromagnets accelerating needle-thin proton beams at 20 trillion eV cryogenic temperature and ultrahigh vacuum. After a decade of work and investment of $2 billion, however, Congress cancelled the program. While much was made in the national press about the wasted effort (and the huge hole in the ground), SSC composites research has yielded unprecedented results that were directly applicable to other ongoing high-energy physics applications (such as the Relativistic Heavy Ion Collider at Brookhaven National Laboratory in Upton, NY, the Tevatron Collider at Fermi National Accelerator Laboratory in Batavia, IL, the Tokomak Physics Experiment in Princeton, NJ, and the European International Thermonuclear Experimental Reactor). For SSC dipole magnets (each nearly 18 288 mm (60 ft) long and weighing 11 340 kg (12.5 tons)) and smaller quadrupole magnets, com-
posite prototype support post tubes were fabricated to test various designs to balance multiple performance parameters with cost constraints (Sondericker 1991; Nicol 1992, 1993; Hiller, 1991, 1992, 1993). Material and process systems examined included continuous carbon fiber/epoxy laminates, injection-molded chopped glass fiber/PEI, and RTMd glass fiber/bismaleimide. Machine-woven near-net shape preforms were also considered, along with RTMd phenolic triazine (McConnell, 1991a) (Fig.45.9). In other components, such as coil spacers and end parts, the predominant prototyping method was machining glass fiber/epoxy, though options considered included highly crosslinked BMI doughs and other polyimide resins with chopped glass, fabricated by lowpressure techniques such as structural RTM. 'Magnets were built for SSC that were never seen before in the world', recounts John Morena, who served as the primary composites materials and processes advisor on the
Fig. 45.9 Glass fiber/epoxy end saddles were fabricated and tested for Superconducting Supercollider magnets. (Source:Brookhaven National Laboratory.)
Applications 979 project. ’There were many “lessons learned” as a result of the years of composite materials development and applications,’ he adds. ’This initial work has provided the composites community with the ability to predict the performance of advanced composites and other polymer materials in cryogenic and high energy physics applications. These original materials are now used in MRI, superconducting magnet energy storage, magnetic levitation systems, fusion, power transmission and space applications.’ Others involved in magnet prototyping were General Dynamics, Westinghouse Electric, Babcock & Wilcox, as well as the Fermi and Brookhaven laboratories and the SSC on-site laboratory. Morena notes a particular materials characterization study conducted as part of Supercollider R&D: subjecting thermoset and thermoplastic resin systems to high proton irradiation at cryogenic temperatures (lo9rads at 4.2K).One intriguing result: undercuring some thermoset composites can increase their longevity in the cold irradiation environment without sacrificing performance (Morena, 1994).
Under the National Maglev Initiative in 1992, four maglev teams developed concept definitions with funding from the Intermodal Surface Transportation Efficiency Act (ISTEA). All concepts were geared toward transport vehicles capable of 482 km/h (300 mph) speeds and incorporated some utilization of composites. Among team members in the early ‘90s were aerospace companies Beech Aircraft and Grumman Corp. Drawing on its extensive aerospace design/fabrication expertise, Grumman Corp. - Aerostructures Div. (Bethpage, NY), began development of a 2/3scale electromagnetic system, based on superconducting iron core magnets spaced 50 mm (2 in) from guideway rails. According to Richard Gran, Director of Advanced Concepts at Grumman, ‘the advantage of this design as compared to a repulsive magnet system is that no secondary suspension system is required; passenger compartment acceleration is monitored (Shaw, 1993).Though Grumman intended to finish construction and testing of its first magnet by April, 1994, funding issues precluded completion: The company merged with Northrop Corp. (Los Angeles) and then with Lockhead Martin by 1997, no maglev 45.2.5 MAGLEV TRANSPORT SYSTEMS OF THE development existed within the newly framed FUTURE business entities. One potential application of SSC research on Also intent upon prototyping a maglev syssuperconducting magnets could be magnetic tem is Maglev 2000 of Florida Corp. (Stuart, levitation (or maglev) of high-speed trains EL.) formed by SSC materials/processes conover elevated guideways. Maglev is consid- sultant and co-inventor John Morena with ered by many to be the transportatiodfreight physicist Dr. John Danby and nuclear engisolution of the future. Composites offer a neer/rocket scientist Dr. James Powell. Danby whole list of attributes in vehicle/guideway and Powell are considered the pioneers of USA structure and in superconducting coils and maglev technology. Prior to publishing their electromagnetic shielding: high strength and first papers on magnetic levitation (Powell, stiffness with low weight, nonconductivity, 1996), Danby reports that concepts of electrononpermeability, corrosion resistance, low dynamic levitation and propulsion had thermal conductivity and high heat capability surfaced periodically but were not considered at low temperature. Overall weight savings of practical. Early in his career, he worked on carbon fiber/epoxy or phenolic vehicle floor building particle accelerators while Powell was beams and shell structure (sandwich panels of involved in advanced reactors. ’Fixed field honeycomb core with carbon fiber/epoxy magnets on a moving vehicle can induce curskins) could better aluminum by 15%of body rents underneath or around the vehicle sufficient to levitate it,’ explains Danby. ’This weight (Cope, 1993).
980 Scientific applications of composites requires a strong force to keep the vehicle separated from the guideway in a stable way. We formulated the geometries to accomplish that strong stability with efficient inductive, shorted metal-loop elements on the guideway and superconducting magnets on the vehicles.’ The highly successful Japanese maglev development is based on the ideas of Powell and Danby (McConnell, 1993a). In 1990, they came together with Morena to build a proposed 32-80 km (20-50 mile) route in Florida with inital design featuring a unique guideway with 8 inch spacing that dampens noise and vibration, and completely enclosing electromagnetic and dynamic fields. It will demonstrate full speed (300 mph) nonmechanical switching capability. Unique to the Maglev 2000 design is the ability to carry and individually deliver heavy freight (such as tractor trailors) to stations on major highways. Specific utilization of composites will likely draw upon Morena’s SSC experience with lowcost fabrication processes such as RTM and resin film infiltration. Guideway and vehicle test articles are being built in 1997. The cost of Phase 1is estimated at $4 million with completion slated for June 1998; in Phase 11, a full scale 120 ft long advanced transport freight and passenger vehicle would be constructed with composite outer secondary shell.
REFERENCES
Blair, Christopher and Petrie, Brian C. 1992. Low earth orbit environmental effects on composite materials: results from LDEF. Intern. SAMPE Tech. Conf. Proc., October 20-22, Toronto, (24): T186-T200. Cope, David B. 1993. Why does maglev need composite materials?Intern. SAMPE Tech. Conf. PYOC., October 26-28, Philadelphia, (25): 141-151. Friebele, Elaine. 1992. The ghost of LDEF-future: EOIM-3 continues the quest. LDEF Newsletter, PO Box 10518, Silver Spring, MD 20914. May 15: 19-22. George, Peter and Dursch, Harry W. 1992. Low earth orbit effects on organic composite materials flown on LDEF. Paper read at NASA-sponsored 2nd LDEF Materials workshop in October, Huntsville, AL. Hiel, Clement, Dittman, Dan and Ishai, Ori. 1993. Designer’s comer: composite sandwich construction with syntactic foam core. Composites 24(5):447450. Hiller, M.W. et al. 1991. SSC quadrupole magnet cryostat design alternatives. Adv. Cryogenic Engineering 37A, New York: Plenum Press, pp. 535-542. Hiller, M.W. and Waynert, J.A. 1992. A cryogenic support post for SSC quadrupole magnets. Proc. IISSC 4, New York Plenum Press, pp. 297-300. Hiller, M.W. et al. 1993. A low-cost support post for SSC quadrupole magnets and other cryogenic applications. Paper read at International Materials Conference July 12-16, in Albuquerque, NM. Kilpatrick, Mark C., Girard, Joseph D. and Dodson, Kelly J. 1992. Design of a precise and stable 45.3 CONCLUSIONS composite telescope structure for the UVCS, April SPIE conference proceedings, Design of The scientific journey continues, surviving Optical Instruments, (1690):196-215. budget cuts, continual goal assessments, and a Lubin, George and Dastin, Sam 1982. Aerospace changing global culture. Composites are Applications of Composites. Chap. 28 of intrinsic to reaching any progressive destinaHandbook of Composites, (ed. George Lubin), tion, whether that be in space or moving New York: Van Nostrand Reinhold, pp. people and goods over this planet. Science722-743. based programs have proven invaluable in Lynch, Ted. 1993. The Technology Reinvestment Project - the government and industry as partgenerating technology transfer for new prodners. Advanced Composites, July/August: 36-42. ucts that enhance essential quality of life, such as medical care. Advanced materials are key McConnell, Vicki P. 1989. Advanced composites make a case in space. Adv. Comp., Sept/Oct: 34enablers, then, for both public and private 45. 1990. National Aerospace Plane: a new enterprise, and for the ultimate science of regime in flight. Adv. Comp., Nov/Dec: 3745. shaping the future. 1991a. Superconducting supercollider: design changes. Composites l n d u s t y Monthly 19, item 3.
References 981 1991b. Progress report: composites in spacecraft. Adv. Comp., July/Aug: 26-34.1992a. Flight vehicles of the future. Adv. Comp., Jan/Feb: 28-34. 1992b. Tough promises from cyanate esters. Adv. Comp., May/June: 28-37. 1993a. Composites ride the rails. Adv. Comp., March/April: 28-37. 199310. Surviving space: a program and materials quandary. HigkPerformance Composites, Nov/Dec: 22-28. 1993c: Focus on design: telescope structure. HigkPerformance Composites, Nov/Dec: 54-56. Morena, John. 1994. Invaluable materials science from Super Collider. High-Performance Composites, Jan/ Feb: 12-1 3. Nicol, Thomas H. 1992. SSC 50mm collider dipole cryostat single tube support post conceptual design and analysis. Proc. IISSC 4, New York: Plenum Press, pp. 747-755.. Nicol, Thomas H. 1993. Single tube support post thermal analysis and test results. Paper read at IISSC 5, May 6-8, in San Francisco, CA. Peterson, Ivars. 1992. Flash tracks-building an eyepiece for a particle accelerator. Science News 142, July 4: 8-9. Powell, James and Danby, Gordon. 1966. High speed transport by magnetically suspended
trains. Paper 66-WA/RR-5 read at ASME Winter Annual Meeting Nov. 27-Dec. 1,in New York City. Shaw, P.1993. Overview of maglev vehicle structural design philosophy, material selection, and manufacturing approach. Paper read at Maglev 93 conference May 19-21, at Argonne National Laboratory. Sondericker, J.H. 1991. Alternate concepts for structurally supporting the cold mass of a superconducting accelerator magnet. Proc. IlSSC 3, New York, Plenum Press, pp. 175-189. Stein, Bland A. 1993. LDEF materials overview. Second Post-Retrieval Symposium, LDEF-69 Months in Space, NASA Conference Publication 3194, (ed. Arlene S. Levine), Vol. 3, pp. 741-789. Tennyson, R.C. and Manuelpillai, G. 1993. Analysis of LDEF micrometeroid/debris data and damage to composite materials. Second Post-Retrieval Symposium, LDEF-69 Months in Space, NASA Conference Publication 3194, (ed. Arlene S. Levine), Vol3, pp. 493-512.
CONSTRUCTION
46
Ever J. Barber0
46.1 INTRODUCTION
Composite materials are used by the construction industry to replace or complement conventional materials such as steel and concrete. The main reasons for the use of composite materials are corrosion resistance, electromagnetic transparency and weight savings. Frequently, structural engineers take advantage of more than one salient feature of composites to formulate a design that is competitive with an alternate design based on conventional materials. Corrosion resistance is the most important advantage of composites with respect to steel for construction applications. The selection of a composite material usually begins with the selection of a resin that is capable of resisting the attack of a corrosive substance. The corrosive agent can be anything from spring water to sulfuric acid. Most composite manufacturers provide corrosion resistance guides for their products. For example, a table listing the maximum operating temperature of isophthalic polyester and vinylester resins as a function of the chemical type and concentration is given by TUFSPAN Technical Data and Design Guide (1991). Chemical resistance of common resins used in pultrusion to various chemical and concentrations as function of operating temperature is given by Pletcher (1991). Fibergrate (1992) supplies a Chemical Resistance Guide for their molded fiberglass
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
and pultruded products listing concentration, operating temperature, and frequency of exposure for a variety of chemicals. Most fiber reinforcements are usually corrosion resistant. Unlike metals, composites do not produce interference with electromagnetic radiation. The resin system can be selected to obtain very low loss factors, but standard resin systems are adequate for most structural applications. Buildings for electromagnetic interference (EMI) testing must be non-magnetic to avoid attenuation and interference with the phenomenon that is being measured. All computer equipment, for example, must be tested in an EM1 facility. Imaging equipment such as nuclear magnetic resonance (NMR) in hospitals must be mounted in a magnetically free environment. An electromagnetically transparent cover for communications equipment allows for the use of less-expensive, non-environmentally protected electronic hardware and reduced maintenance costs. Antennae structures that do not interfere with the signals being relayed or received by the antennae increase the efficiency of the system. While weight saving is the main driving force behind the application of composites in aerospace, it is not so critical in construction projects. However, reduction of structural weight can be exploited as a secondary advantage to help offset the higher cost of composites as compared to conventional materials. Lightweight structures require less foundation and supporting structure. In the case of bridges, a noncorrosive bridge deck can be built to replace existing steel reinforced
Current applications 983 concrete decks that corrode rapidly under the attack from de-icing chemicals. For example, a pultruded deck was used to construct the Wick Wire Run vehicular bridge on public road 26, in Taylor County, West Virginia (completed August 1996).An added advantage of a composite deck would be the weight reduction that supposedly would allow the user (highway department) to re-rate some bridges for a higher live load without major modifications to the existing superstructure. The live load could be increased by approximately the same amount of dead load saved with the use of the composite deck minus adjustments for dynamic effects. Other applications where weight savings are important are cladding of buildings, rehabilitation of chimneys, etc. 46.2 CURRENT APPLICATIONS
Current applications of composites in construction can be classified by the major advantage of the composite material that is exploited. The main ones are corrosion resistance and magnetic transparency. An alternative classification may be based on the type of construction. Composites are used in the form of structural shapes (similar to steel construction), as reinforcement for concrete, cables, and for rehabilitation of existing structures. Reinforcement of concrete may be in the form of conventional reinforcement, prestressed concrete, or post-tensioned structures. Rehabilitation applications include repairing deteriorated structures as well as increasing the load carrying capacity of sound structures to re-rate them for higher load capacity. Some of the recent applications of composites in construction of civil infrastructure will be described in the next section. Further examples of applications can be found in previous reviews (Barbero and GangaRao, 1991, GangaRao and Barbero, 1991), professional journals ( e g SAMPE Journal, ASCE Journal) and edited books (Mufti, Erki and Jaeger, 1991a, 1991b; Iyer, 1991, Neale and Labossiere, 1992).Applications to pipes and tanks in ser-
vice for more than twenty years atest the excellent corrosion resistance of fiberglass reinforced composites. The following examples of fiberglass reinforced isopolyester resin applications, described extensively in the excellent review by Adams and Bogner (1993), illustrate the feasibility of constructing composite structures and using them for many years. A three mile pipe of diameter 254 mm (10 in) is reported in service since 1971. Fifteen miles of piping, carrying saline water with temperature up to 50°C (112"F),pressure up to lo6 N/m2 (10 bar), and exposed to sunlight has been in operation in Saudi Arabia since 1983 without problems. An old sewage duct was lined in 1971, then inspected in 1991 showing no sign of deterioration. More than 300 000 underground fuel tanks are in use in North America alone. Some of these have been in operation for more than 26 years without problems. Internal or external lining of steel tanks has been common practice for more than twenty years in the oil industry to protect and reinforce the bottom of steel tanks that are corroded internally because of corrosive substances in the oil or externally because of contact with soil. Wine tanks have been in operation for more than 20 years without problems. Ducts carrying chlorine gases and sulphur dioxide, in use since 1962, and tanks holding hydrochloric acid, in use since 1964, remain in perfect condition. A chimney exposed to organic chemicals, water vapor, and temperature up to 60"C, in operation since 1968 is reported. Chimneys are important construction applications because they are load carrying structures designed for large wind loads. A detailed account of recent applications of composites in construction, classified by the type of construction, follows. 46.2.1 REINFORCEMENT OF CONCRETE
Concrete can be reinforced with fiber reinforced composites, with fibers mixed in the concrete, or by polymers added to the concrete mix. This article will concentrate on the use of
984 Construction fiber reinforced composites, i.e. a combination of fiber and polymer matrix, to substitute or complement the traditional use of steel reinforcing bars (rebars) in concrete. The addition of polymer or fibers to cementitious materials, sometimes better classified as a ceramics, bearing limited resemblance to regular, low cost concrete, falls beyond the scope of this article. The interested reader may wish to consult Mufty, Erki, and Jaeger (1991b, Chap. 2). Use of fiberglass and carbon fibers mixed directly into concrete has been reported (Mufty, Erki and Jaeger, 1991a).Seibu Construction Co. Ltd. used carbon fibers, produced by Mitsubishi Kasei Co., for the exterior walls of the Kitakyushu Prince Hotel in Japan. Kajima Co. also used carbon fiber reinforced concrete panels for the exterior of its head office building in Japan. Polymer concrete is used in a variety of applications, such as in highway parapet walls developed by Morrison Molded Fiber Glass (Fig. 46.1). The installation of the lightweight polymer concrete panels reduces to anchoring the panels to an existing parapet, then pouring concrete into the panel that act as a stay-inplace concrete form. Composite reinforcement of regular, low cost concrete can be done using rebars, grids, pre-stressing tendons and posttensioning cables. F
Rebars Reinforcing bars (rebars) have been used for reinforcing concrete structures that require magnetic transparency, such as imaging equipment at hospitals (Fig. 46.2). The major expected application is to replace steel rebars in concrete bridge decks exposed to de-icing chemicals. The first use of composite rebars for a vehicular, public bridge is the McKinleyville Bridge, located in West Virginia, USA. Construction was completed August 1, 1996, by Orders Construction Co., of St. Albans, WV. The bridge is a 54 m (177 ft) long concrete deck over steel stringer accommodating two lanes of vehicular traffic. The reinforcement of the concrete is exclusively made of composite rebars with E-glass fibers. It was designed by the West Virginia Department of Transportation - Division of Highways (project S305-27/4-0.03) and the Constructed Facilities Center at West Virginia University. The project was supported by the Federal Highway Administration, composite rebar
1
Fig. 46.1 Installation of polymer concrete panels as highway parapets. (Courtesy of the Quazite Division o~MMFG.)
Fig. 46.2 Fiberglass reinforced composite rebars during construction of reinforced concrete structure. (Courtesyof Reynolds, Schlattner,Chetter, and Roll Inc.)
Current applications 985 manufacturers, and the Corps of Engineers. E-glass polyester rods are cost competitive with steel rebars at the present time. These rebars have evolved from smooth pultruded rods to engineered rebars with improved bond strength to concrete. Further improvements in bond strength, tensile strength, and durability are certain to occur in the near future. E-glass composite tanks have been in service for over 20 years, in permanent contact with highly acid environments and under constant stress. The use of the SPI liner (ASTM D-3299 and D-4097) proved very successful to protect glass fibers from attack by chemicals in storage tanks. Also, composites have been in contact with concrete for many years without deterioration. Investigation of the possible degradation of fiberglass rebars in concrete is under way. The possibility of moisture intake of Aramid fibers may be a problem for rebars expected to be in service for at least 50 years. Carbon fibers have superior properties but their cost limits their potential as a replacement for construction steel. Rebars are placed in the form-work to provide for reinforcement of concrete in the same way as fibers are used to reinforce polymers. Usually, two perpendicular orientations are used, with more rebar area in one direction, according to the requirements of the structure. Placement of rebars requires relatively inexpensive labor. The grid-like structure necessary to reinforce concrete can be obtained preassembled in the form of grids. Grids are rectangular networks of rebars produced with continuous fibers. Although more expensive than rebars, they may be convenient if the labor cost is high or the installation difficult, as in the case of tunnels. Grids can also be produced at or near the construction site by tying rebars into a grid. The light weight of the resulting grid makes transportation to the site and installation simple.
sion strength and, for most practical purposes, negligible tensile strength it is advantageous to pre-stress concrete so that a state of compressive stress is created before the actual load is applied. (The American Concrete Institute recommendation ACI 363 R suggests a value for the modulus of rupture of concrete f, = 11.7 cfc)1/2which is a very low value of tensile strength in relation to the compression strength of concrete, f,.)Then, applied loads only reduce the amount of compressive stress without producing tensile stresses. In this way, concrete cracks are prevented, which in turn reduces moisture intake and degradation. High strength steel tendons are currently used for pre-stressing concrete. Even though cracks are arrested by the prestress, concrete is porous and water and chemicals may reach the prestressing tendons. Composite tendons may replace the steel tendons for added durability. Pre-stressed concrete is usually pre-cast at a factory, then transported to the site. The tendons are prestressed by a hydraulic jack, the concrete is poured and left to cure. After the concrete is set, the tendons are cut to remove the chucks used to apply the pre-stressing load. Some of the pre-stress is lost because of the compression, creep, and shrinkage of concrete. The higher the elastic modulus of the tendon, the higher the pre-stress loss. For this reason, low modulus glass fiber tendons experience less pre-stress loss than steel or carbon fiber tendons. Pre-stress loads are high and induce large strains in the tendons, which may accelerate degradation if the tendons are exposed to highly alkaline environment (Sen, 1992). Two important requirements for pre-stressing applications are a good bond strength between the tendon and the concrete and availability of an effective temporary anchorage system. The anchor is not very critical because it is temporary. Therefore, it does not have requirements of non-corrosiveness, cost, size, etc., but it must be able to transfer the Prestressing load and sustain it until the concrete sets. Since concrete has a relatively high compres- Anchorage systems are reported by Noritake
986 Construction
et aZ. (1992), Mochida, Tanaka and Yagi (1992), Kakihara et al. (1991), etc. for various types of tendons based on glass, aramid, and carbon reinforcement. Sen (1992) used the anchorage developed by Iyer at South Dakota School of Mines and Technology. Efficient utilization of composite tendons calls for large pre-stressing forces to be applied. Permanent levels of stress in the tendons should be under the stress rupture limit of the composite (see Section 46.3.4). The pre-stress forces applied to composite tendons may induce cracking of the polymer matrix. Matrix cracking;may be detrimental to the fatigue life of the composite. Also, cracking facilitates the ingress of moisture that may cause degradation of organic fibers or precipitate alkaline reaction of glass (see Section 46.3.4). Pre-stressing has been successfully used by Iyer (1993) in a bridge deck (Fig. 46.3) precast and pre-stressed, then transported to the bridge site. The bridge construction was sponsored by Owens-Corning, AMOCO, South Dakota Office of Economic Development and South Dakota Cement Plant; with the participation of Clark Engineering, Glenn C. Barber and Associates, FMG Engineering, Polygon, Shell, Central Mix, Carlon/Ace Hardware and South Dakota Concrete Products. A 180 mm (7 in) concrete deck was post-tensioned with glass and carbon fiber tendons which were subsequently grouted with epoxy based mortar. The deck is supported by steel girders 2.6 m (8.5 ft) apart and the bridge spans 9 m (30 ft). Aramid bars were used in a standard prestress pre-cast concrete factory to build concrete barges (Noritake et aZ., 1992). At least 10 concrete bridges have been constructed in Japan with some kind of composite reinforcement (Mufti, Erki and Jaeger, 1991a). The Shinmiya bridge is a pre-tensioned concrete bridge built in 1988 in a coastal area. The precast concrete girders were pre-stressed with seven strand, carbon composite tendons produced by Tokyo Rope Manufacturing Co. Epoxy coated steel rebars were used for the "
-
-'*----
I
J.
I I
r,_
d
Fig. 46.3 Post-tensioned concrete-deck bridge over steel stringers. (Courtesyof Owens Coming.)
stirrups. The Sumimoto Construction Co. built a demonstration bridge in 1990 in Oyama using three concrete box girders pre-tensioned with aramid composite rods, produced by Teijin Co. Grouted anchorages were used to pre-stress the rods at a permanent stress of 70% of their static tensile strength. While the stirrups were also aramid composite rods, epoxy coated steel rebars were used as shear connectors between the girders and the reinforced concrete deck, which was reinforced with epoxy coated steel rebars. The deck was post-tensioned with aramid composite rods which were then permanently grouted in the deck. The Birdie bridge was built by Kajima Co. in 1990 at the Ibaragi Prefecture using a variety of composite materials. Carbon fiber composite rods, produced by Mitsubishi Kasei, were used to anchor the abutments. Precast concrete panels reinforced with vinylon short fibers are connected with a grid of carbon fiber composite cables produced by Tokyo Rope Manufacturing Co. The bridge is prestressed with aramid composite flat bars commercialized by Nippon Aramid Limited. Half of the aramid bars are used as pre-tensioning tendons and the remaining act as post-tensioning tendons,with a permanent
Current applications 987 stress of one third the static tensile strength. Post-tensioning Post-tensioning of concrete with steel or composite tendons is performed to induce a state of compressive stress in the concrete similarly to pre-stressing, but post-tensioning is performed at the construction site. A hole or some kind of access is left in the concrete to thread the post-tensioning cables through. Tensile force is applied to the tendons against the concrete structure before the structure is loaded by the service loads. The anchorage system is subject to severe requirements. First, it must sustain the tension load for the whole life of the structure, which requires very careful design against creep in the anchor. Second, the anchor should be resistant to corrosion as the tendon itself if the non-corrosive properties of the system are the objective of replacing steel tendons. The reduced weight of the tendons is not very important because the heavy weight of the concrete structure. While the stress losses of fiberglass tendons are smaller than those of steel cables, it is difficult to justify the use of more expensive and novel composite tendons for this reason alone. Therefore, corrosion resistance is the main objective of using composite tendons. In this case the anchorages should be resistant to corrosion. Also, the structure to be posttensioned is usually made of reinforced concrete (although the reinforcement is not sufficient to carry all the load). If corrosion is a problem, the conventional reinforcement (not pre-stressed or post-tensioned) may also have to be made of composite material rebars. Anchors for post-tension applications are still being developed and evaluated. Meisseler and Preis (1989) report on an anchor developed to hold glass reinforced tendons. The anchor is of the potted type, in which the tendons comprising a cable are spread at each end and potted in a steel anchor with some grouting material, usually a polymer. The potted type anchors are based
on the transfer of the axial load in the cable, by shear in the grouting material, to the anchor. Attaching the anchor to the tendons is a labor intensive process. The resin used to bond the tendons to the anchor may creep, leading to loss of pre-stress. This anchor was used to partially post-tension several bridges. Cases have been reported where individual tendons slipped and broke at the anchor during post-tensioning at the bridge site. Porter and Barnes (1991) report several anchorage systems. To avoid the problems that all potted type anchors have, Ahmed and Plecnik (1989) developed a filament wound cable where continuous fibers are wound around the end eyes of the cable. Each cable must be custom made for the required length, but the problems of potted type anchors are eliminated. The Bachigawa-Minami-Bashi bridge (Koga et al., 1992) uses both pre-stressing and posttensioning carbon fiber tendons. The Schiessbergstrasse bridge in Germany and the Notch bridge in Austria (Wolf and Miesseler, 1992), use partial post-tensioning with fiberglass cables. Some of the tendons that form a cable contain sensors (copper wire or fiber optical gauges) to monitor the strain level, integrity of the tendon, and location of eventual damage. A demonstration bridge was built in 1990 by the Sumimoto Construction Co. in Oyama, Japan, using a single concrete box girder, post-tensioned with aramid composite rods produced by Teijin Co. The internal tendons are placed in a parabolic housing in each web and post-tensioned with steel grouted anchorages to a permanent stress of 25% of the static tensile strength of the tendons. The external tendons are placed at the bottom of the box girder and post-tensioned, with grouted anchorages built with composite casings for added corrosion protection, to a permanent stress of 10% of the static strength. Besides the composite post-tension cables, the regular reinforcement of the girder uses epoxy coated rebars. Further details are given by Mufti, Erki and Jaeger (1991a).
988 Construction
Rehabilitation There is significant interest in using composite materials for rehabilitation and upgrading of existing structures. These structures may have been damaged as a result of corrosion of the steel reinforcement or they may need upgrading to new seismic standards, larger traffic loads and volume, etc. Interest in composites is motivated by the ease of installation of the reinforcement, its corrosion resistance (mainly at the bond surface), and the possibility of selecting from a variety of elastic modulii which improves the compatibility between the reinforcement and the existing structure Meier et al. (1992), Saadatmanesh and Ehsani (1991), and others have demonstrated the feasibility of rehabilitating concrete, steel, and wood structures by reinforcement with composite plates. Meier et al. report a variety of failure modes that may be encountered in concrete beams reinforced with composite plates. The reinforcement may be a cured composite plate bonded to the beam (Munipalle, 1992), or a room-temperature cure prepreg directly applied to the structure. The reinforced plates can be applied with or without pre-stressing. While pre-stressing the composite plate increases the efficiency of the reinforcement, it also complicates the rehabilitation process. Since delamination of the reinforcing element is of major concern, special reinforcement details are used at the ends of the reinforcement. Composite materials can also be bonded to the sides of beams to improve the shear strength. While using prepreg materials, both shear strengthening and resistance to delaminations can be obtained by partially or completely wrapping the beam with the reinforcement. Meier et al. (1992) reports on the rehabilitation of two bridges in Switzerland. The Ibach bridge was repaired using carbon fiber reinforced plates bonded to the concrete bridge. The rehabilitation of the timber bridge in Sins involved replacement of the wood pavement with transversely pre-
stressed and glued laminated deck (Davalos and Salim, 1992). Some of the timber beams were reinforced with carbon reinforced epoxy plates. Rostasy, Hankers and Ranish (1992) report on the reinforcement of the Kattenbusch bridge (Germany) using steel and glass reinforced composite plates. The main reason cited for the selection of composite plates is the documented corrosion of the steel plates at the bonded interface. Composite jackets have been proposed to retrofit columns for seismic solicitations. Scale samples (40%) were tested (Priestley, Seible and Fyfe, 1992) and promising results were found. A fiberglass epoxy jacket is built around the hinge region of the column. On an active confinement jacket, hoop stress is induced by pressure grouting the space between the concrete column and the composite jacket with either epoxy or concrete. A passive confinement is obtained when no pressure grouting is used. Passive confinement is used in regions of high compression stress, over the region of lap-spliced longitudinal bars, or for regions of high shear stress. Significant concrete dilation and consequent longitudinal microcracking is necessary to activate the confinement effect in passively encased concrete. Active confinement has the advantage that the confinement effect is always available. Chimneys have been rehabilitated in Japan by Mitsubishi Kasei using carbon composite tape and strands (Mufti, Erki and Jaeger, 1991a). The composite tape is applied along the length of the chimney to provide additional bending strength. The strand is wound around the chimney to provide hoop reinforcement. A fire protection mortar is finally applied to protect the polymer composite from fire, to limit the heat gain, and to avoid degradation of the polymer matrix that otherwise would occur if the carbon composite were exposed to the environment.
Current applications 989 46.2.2 STRUCTURAL SHAPES
Platforms
Composite structural shapes resemble steel hot- and cold-rolled structural sections. Composite shapes are produced by pultrusion in a variety of sizes and shapes (Creative Pultrusions, 1989; MMFG, 1992). The reinforcement of choice is E-glass mainly because of low cost. Most shapes are produced with isophthalic polyester and vinylester, but some pultrusions using epoxy, phenolic, and even thermoplastics are available at higher cost. Even for the most common polyester and vinylester, there is a wide variety of resins systems, differentiated by mechanical properties, thermal behavior, and cost. While vinylesters are considered to have superior corrosion resistance and mechanical properties than polyesters, high grade polyesters may match the mechanical properties of vinylesters. Structural shapes are used primarily because of their excellent resistance to chemical degradation. In the case of electromagnetically transparent structures, structural shapes are routinely used for building construction in very much the same way as steel shapes. However, E-glass reinforced shapes have a much lower modulus of elasticity than steel, which causes design problems when direct replacement of steel shapes by composite shapes is attempted. Specialized sections, with cross sections different from steel shapes, are available for use in building systems with construction characteristics (e.g. joist spacing) very similar to steel frame buildings (Composite Technology, 1992). Unlike steel shapes, the mechanical properties of composite structural shapes largely depend on the internal reinforcement, and the thermal and corrosion response varies drastically with the resin system used. Since the reinforcement and the resin systems used are not standardized, very different set of properties are possible for identical cross sections, from different manufacturers.
Industrial platforms built entirely of pultruded fiberglass reinforced plastics are widely used because of their corrosion resistance. An area of application of special interest is offshore platforms for oil production. Two views, from above and below, of a fiberglass well bay platform are shown in Figs 46.4(a) and (b). Steel well bay platforms corrode quickly in the marine environment. Painting jobs are difficult because sand blasting may release paint into the ocean unless costly precautions, such as scaffolding, are taken. Old paint may be lead-based and cannot be stripped unless it is completely captured and dumped at a hazardous waste site. Well bay platforms are installed after the production wells are in operation. The area is usually cluttered with equipment, which makes difficult the access of heavy lifting equipment needed for installing heavy steel platforms, and welding cannot be done without shutting down oil production. Fiberglass platforms weigh typically 30% of their steel counterpart, allowing for installation by a smaller crew, in less time, with less demand for space for lifting equipment, and reduced transportation costs to the site. The composite platforms are assembled with mechanical connections eliminating the need for welding and they are virtually maintenance free in the corrosive marine environment, without need of painting. A 6 m x 12 m (20 ft x 40 ft) well bay platform was installed in 1986 on Shell’s Southpass 62 production platform in the Gulf of Mexico by a crew of four in two days and it is estimated that a similar steel platform would have required five days by a crew of eight because of the lack of room for heavy lift equipment.
Building systems Complete building systems, including the structural frame, cladding and roofing are available in fiberglass reinforced composites (Composite Technology, 1992).Electromagnetic
990 Construction
Fig. 46.4 Well bay platform in an off-shore oil production facility (a) seen from above, @) seen from below. Composite materials manufactured by MMFG. (Courtesy of MMFG.)
Current applications 991 Interference ( E M ) testing buildings must be built of materials free of magnetic interference, with all metals ruled out and the use of timber limited by the need to use steel connectors. Pultruded structural shapes, fiberglass cladding, and foam core panels offer an attractive alternative since all components can be connected with fiberglass bolts and or glued together to form an electromagnetically transparent structure. Innovative combinations of glued laminated timber (GLULAM)with composites have also been exploited. A computer testing facility (Fig. 46.5) was built for IBM in Poughkeepsie, NY, by Corflex International Inc. of Warren, OH, and Haines Lundberg Waehler of New York, NY, using pultruded structural shapes from Creative Pultrusions Inc. (1989).
Fig. 46.6 Fiberglass spire installed atop the 55-story Nations Bank in Atlanta, GA. Comuosite materials manufactured by MMFG. (Courtes; of MMFG.) Fig. 46.5 Computer testing facility using pultruded structural shapes and fiberglass reinforced panels. (Courtesy of Creative Pultrusions, Inc.)
The Nations Bank building in Atlanta, GA, shown in Fig. 46.6 features a 11m (36 ft) tall, all fiberglass spire at its top. Electromagnetic transparency of the composite material used allows the spire to house valuable communication antennae. Light weight, molded-in color, and timely delivery were cited as advantages of composite materials for this highly visible application. The spire is mounted 312m (1023ft) above ground, and it is designed to sustain a wind pressure of 550 kPa (BO psi).
Pedestrian bridges A number of pedestrian bridges have been built with composite materials. Early bridges are reported by Meier (1991), Barber0 and GangaRao (1991) and others. While a few bridges have been built by hand lay-up, the majority are built with pultruded sections. Concrete bridges partially reinforced with composite tendons are discussed in previous sections on concrete reinforcement. In this section, we concentrate on bridges where most of the materials used in the construction of the bridges are composites. Because of cost, E-glass reinforcement has been used most. Johansen et al., from E. T. Techtonics of Philadelphia, PA, report on the construction of
992 Construction
shapes. (Courtesyof E.T. Tcchtonics.)
three bridges of 6.1 m (20 ft), 9.75 m (32 ft), and 15.24m (50 ft) (Fig. 46.7) using their design and construction method, called PRESTEK. Given the low stiffness of E-glass reinforced pultruded composites when compared to steel, conventional steel or concrete designs are not efficient when implemented with composites. The PRESTEK system uses a beam-truss geometry of pultruded tubes prestressed with aramid or steel cables. The three bridges are a king-post truss, a queen-post truss and a bow-string truss respectively. Maunsell Structural Plastics in Beckenham, Kent, England, with the participation of GEC Reinforced Plastics (pultrusion manufacturing), Scott Bader Co. Ltd (resin supplier), Vetrotex UK (glass reinforcement), Ciba Geigy Plastics (adhesives), Linear Composites (Parafil cable stays), R. O'Rourke and Sons (construction management), and University of Dundee, constructed a pedestrian bridge over the river Tay in Aberfeldy, Scotland. The pultruded deck is cable stayed from two Aframe towers, 17.5m (57 ft) high, to provide for adequate stiffness despite the low stiffness
of the pultruded glass-polyester deck (Fig. 46.8). The bridge is 113m (370ft) long and 2.23 m (7.3 ft) wide, with a main span of 63 m (206 ft). The deck and towers are constructed with Maunsell's interlocking panels (ACCS System by Designer Composites Technology Ltd), pultruded with 70% volume of fiberglass reinforcement and isophthalic polyester resin, then bonded together with an epoxy adhesive. The cable stays are Parafil ropes, a Kevlar fiber core in a polyethylene sheath. The bridge deck was designed to be modular so that every component could be handled by hand. The heaviest module, a 6 m (19.7 ft) by 0.6 m (2 ft) plank weighed only 66 kg (145 lb). The weight of the deck is only 150kg/m (lOOlb/ft) and each tower weighs only 2500 kg (5511 lb), facilitating erection and reducing the cost of the foundation. The composite system has class 2 fire resistance rating. Bridge enclosures
The A19 Tees Viaduct in Middlesbrough, UK, is a steel-concrete bridge with 117 m (383 ft) span
Current applications 993
Fig. 46.8 Pedestrian bridge across the river Tay in Aberfeldy, Scotland, features an all-composite deck cable stayed with Parafil cables from two composite A-frame towers. (Courtesyof Maunsell Structural Plastics Ltd.)
(Head, 1988) where rapid deterioration of the steel plate girders was taking place. Maintenance (e.g. painting) and rehabilitation work are difficult since the viaduct spans over railroad tracks. A composite material enclosure was built in 1989 to prevent further deterioration by isolating the bridge from the environment and to facilitate maintenance and rehabilitation. A floor area of 16000m2 (172 200 ft2)using 250 metric tons (275 US tons) of composite material was created under the bridge by interlocking pultruded panels wrapped around the steel girders of the bridge to create a box. The enclosure system was designed by Maunsell StructuralPlastics Ltd, of Beckenham, Kent, UK. The system uses interlocking panels designed by Maunsell and fabricated by GEC Reinforced Plastics Ltd, while the main contractor was Fairclough Construction Ltd. The enclosure system has been also used in the construction in 1992 of the enclosure to the Bromley South Bridge, seen in
Fig. 46.9, over the railway station. Pultruded plank and connector sections are joined to form a floor system suspended from the steel girders of the bridge. Pleasant appearance was required since the bridge is located in a residential area and over the railway station. Light weight and low maintenance costs were cited as additional advantages of composites for this application.
Cooling towers The resistance of composite materials to humidity and creep under sustained loads has been demonstrated by their successful application to cooling towers, in operation for more than twelve years (Fig. 46.10). Cooling towers are permanently loaded with the heavy weight of the ceramic filling used for cooling. These towers are built by Ceramic Cooling Towers of Forth Worth, TX, entirely of composites except for the ceramic filling. The
994 Construction
Fig. 46.9 Enclosure of the Bromley South bridge is accomplished with interlocking pultruded panels bonded together into a floor system and suspended from the steel girders of the bridge. (Courtesy of Maunsell Structural Plastics Ltd.)
composite material is in permanent contact with hot water and humid hot air circulating through the tower. Reduced maintenance of the tower is cited as an important advantage of using composites. Light weight of the structure of the tower is an additional advantage when the towers must be located on top of buildings. Construction of large industrial cooling towers motivated the development of the UNILITE Modular System which takes advantage of modular construction to reduce cost. Composites facilitate modular construction because of the variety of shapes that can be easily produced. Marine construction
Potential applications for composites in the construction and rehabilitation of marine and waterfront constructed facilities include docks, piers, harbors, etc. An all-composite pontoon designed by Maunsell Structural Plastics Ltd, of Beckenham, Kent, UK and dis-
tributed by Designer Composites Technology Ltd, UK, is shown in Fig. 46.11. Finger piers of up to 15 m (49 ft) are built with cellular interlocking pultruded panels, joined to foam filled flotation units. The surface of the pontoon is covered with a non-slip polyurethane coating which is acid, solvent, and heat resistant. 46.3 DESIGN CONSIDERATIONS
Current composite design practice emphasizes simultaneous design of the structure (beams, plates, frames, etc.) and the material (composite) for optimum performance. Design practice for composites used in construction differs from aerospace applications in the sense that standardization of components is required because of cost and safety. Therefore, design is divided into member design, usually performed by the manufacturer, and system design, carried out by the structural engineer. Codes of practice do not yet exist for structural design with composites, but they are being
Design considerations 995 . ..
...
._...
.
deflections are controlled by the bending stiffness D, equivalent to EI for steel beams. Since composite beams have different values for modulus of elasticity at various points in the cross section, it is not possible to completely define the stiffness with respect to both axis of bending by the product of the modulus and the corresponding moment of inertia (€1 and €Ixx ) Instead, two bending stiffnesses and Dx are defined, with respect to the strong axis and weak axis respectively. Shear deformations are neglected for steel beams because the shear stiffness of steel is high ( G / E = 0.4) while for composites is low ( G / E < 0.1). The shear stiffnesses of a composite beam with respect to the two axis of bending are denoted F y and Fx. The values of the bending, shear, and axial stiffness (Dy, K F ~ A,) , can be obtained from tables supplied by the beam manufacturer, k obtained experimentally (Bank, 1989), or computed if detailed information about the constituent materials is known (Barbero, 1998). Maximum deflections can be computed using the formulae in Fig. 46.12. Experimental values of shear stiffness usually refer to the Fig. 46.10 All-composite cooling towers like this, product of K F where ~ K is the shear correction developed and manufactured by Ceramic Cooling factor. Tower, have been in operation for more than twelve years.
<
developed. Some methods of analysis and design recently developed are described in this section.
c
46.3.1 BEAMS
Beams are the most common structural component in civil engineering applications. Both deflection and strength are equally important in the design of composite beams. Composite beams are thin-walled and composed of an assembly of flat panels. Most beams are prismatic but they can have taper. They are produced by pultrusion, filament winding, hand lay-up, automated lay-up, etc. Deflection of composite beams has two components, bending and shear. Bending
. -+-----
Fig. 46.11 The pontoon pier shown is built of interlocking pultruded panels and foam filled flotation devices. A polyurethane coatingprotects the surface from acid, solvents, and heat. (Courtesy of Maunsell Structural Plastics Ltd.)
996 Construction W
5
n
4 W L
s%=384%
+
1 WLZ 8 KF~
I
Fig. 46.12 Center deflection of composite beams including bending and shear effects.
Design considerations The sigruficance of the shear deflection with respect to the bending deflection varies with the span, the larger the span the lesser the influence of shear. Sometimes, properties of beams are reported without distinction between the bending and shear components, using an apparent value of modulus Eapp= Dy/lyy,where D,, is the bending stiffness with respect to the strong axis. The bending stiffness with respect to the weak axis Dx cannot be accurately obtained as Dx = EapIxx. The apparent stiffness Eappis then used in &e classical deflection formulas (for steel) that do not account for shear deformations. The reported values are usually based on three-point bending tests, performed at the factory with a specific span, which is seldom reported. The results of using the classical (steel) deflection formulas for spans or loading other than that of the test are only approximate. The main modes of failure of beams in bending are: (a) compression crushing of the compression flange; (b) local buckling of the compression flange; (c) tensile rupture of the tension flange; (d) shear failure of the web; and (e) web buckling. Since each part (panel) of the cross section can be built with different materials, the failure mode can be controlled by design. Local buckling modes can be eliminated by increasing the thickness and choosing the fiber orientations properly (Barber0 and Raftoyiannis, 1993). The compression strength of composites is lower than the tensile strength. Therefore, a symmetric section is not the most efficient cross section. Symmetric sections fail in the compression flange first. Tensile failure may occur in unsymmetric sections, when the compression flange works with composite action with a deck, etc. Shear failure is less likely to occur in sections with multiple webs. Incorporation of off-axis fibers (cloth, mats, etc.) increases the shear strength. 46.3.2 COLUMNS
Columns are structural members subjected primarily to compression forces acting along
997
the length of the member. Lateral forces (e.g. wind forces) and bending moments (e.g. eccentric loading) are considered secondary forces, which are dealt with separately. Column performance is limited by one of two failure mechanisms, crushing and buckling. Crushing is the failure of the material because of excessive compressive stress, similar to yield of steel. Buckling is more frequent in composite columns because thin-walled sections are preferred. A thin-walled section may experience at least three different types of buckling, which are described next. Long and slender columns fail in a global sense when the axial load reaches a critical value P,, For load values lower that the critical load the column remains straight. When the load reaches critical value, the column experiences sudden lateral deflection. The axial stiffness after buckling is much lower than the stiffness before buckling. Therefore, the lateral deflections are quite large and they usually precipitate another mode of failure like crushing of parts of the cross section, leading to collapse. The concept of slenderness allows us to compare members of different cross sections and column lengths for their tendency to buckle. For a composite column, the slenderness is defined as =
L&)
(46.1)
where P, is the local buckling load, D is the bending stiffness, and Le is the effective length of the column, which is used to account for different end conditions (Gere and Timoshenko, 1990, p. 589). Short and stubby columns, which have a low Slenderness value, are less likely to buckle in a global mode as described previously. However, individual parts (flange or web) of the cross section may buckle locally. Local buckling is very likely to occur in composite columns because they are commonly thin walled. The compressive stress required to trigger local buckling increases with the thickness
998 Construction of the cross section and the local stiffness of the material. The narrower the flange or web under consideration, the higher the local buckling load P,. The wavelength of local buckling is independent of the length of the column for columns of practical lengths. Flange buckling of open section members is triggered by pure compression.It occurs during compression of columns and on the compression flange of beams in bending. Web buckling is initiated by shear and it occurs during bending of beams. Web buckling of open section columns is not common because most open sections have wide flanges that buckle first. There is of course no distinction between flanges and webs in closed section columns. Unlike steel structural shapes, composite closed sections are easier to produce and structurally more efficient than open section members. Local buckling can be prevented by choosing the section geometry and material properties. Global buckling can be prevented by a combination of section geometry, material properties and bracing. If all buckling modes are prevented, the strength of the member is limited by the crushing strength of the material itself, which plays a role similar to the yield strength of metals. The crushing strength is a material property which is independent of the thickness of the flange or web and the geometry of the section. As for any composite property, it depends of the constituents (fiber and resin) and the arrangements of the fibers inside the material (orientation, fiber volume fraction, stacking sequence). The crushing strength is usually determined experimentally but the main factors that influence its value can be highlighted by predictive equations (Tomblin, 1994; Barbero, 1998). A slender column buckles in a global (Euler) mode. A not-so-slender column may fail in a local buckling mode. A thick-walled stubby column may fail due to crushing. Columns with a slenderness ratio less than 0.5 fail in a local buckling mode. For slenderness larger than 1.5 the mode is purely global (Euler). Columns with slenderness between
0.5 and 1.5 show some type of interactive phenomenon. The interaction occurs between the local mode, the global mode, and crushing. Interaction results in lower buckling loads than those predicted by any of the modes acting alone. There are many situations of practical interest for which the buckling loads required to produce two or more failure modes (Euler, local, crushing) may be very close. In this case the failure modes interact. That is, the proximity of the stresses to more than one mode of failure causes the structure to fail at a lower stress value that predicted by either of the modes involved should they be acting separately. Euler and local modes interact to give an overall strength deterioration. Interaction must be taken into account because the strength values predicted by either isolated mode fit are not conservative. The failure load of a column (Po) taking into account local, global, and interaction phenomena simultaneously can be obtained from the following design equation (Barbero and Tomblin, 1993)
The column properties needed to use this design equation are: the local buckling load PL; the interaction constant c; and the bending stiffness D, which along with the length of the column enter in the computation of the slenderness 1 (46.1). All these properties can be determined experimentally or predicted analytically. The bending stiffness D and the local buckling load P,, can be predicted analytically, while an analytical study of the interaction phenomenon is presented by Raftoyiannis (1993). The design equation does not have any safety factor included. Typical properties of wide-flange pultruded structural shapes are given in Table 46.1 where the interaction constant is c = 0.84 for all sections reported.The length at which maximum interaction occurs is denoted by L‘.
Design considerations 999 Table 46.1 Column properties
obtain the Euler buckling load P,,, as the inverse of the slope in the A/€' vi: A plot, Section PL (kN) D (kN cm2) L x (cm) where A is the lateral deflection and P is the load. 102 x 102 x 6.4 223.25 6094.67 105.9 Interaction testing can be performed with 152 x 152 x 6.4 20954.42 221.5 175.12 the same setup described for global (Euler) testing. The objective is to determine the interaction constant c in (46.2). Once two or more modes of failure interact, the Southwell method cannot be used. Therefore, only the Testing of short columns is performed to iden- maximum collapse load is reported for tests tify the local buckling load P, in the column performed on columns having slenderness design (46.2).Buckling of the flanges is seldom values between 0.5 and 1.5. The collapse load a sudden phenomenon as described by the is lower than the critical load that would occur theory because of the imperfections of the should any of the modes involved act isolated material. Therefore, flange lateral deflections from the others, as the experimental data are observed from the onset of the test and clearly indicates. The test is conducted with a they grow as the load is applied. The buckling column length that exhibits maximum interacload P,, to be used in the design corresponds tion (L*), which occurs for a column to the asymptote of the hyperbolic curve of slenderness 1= 1. The interaction constant is load vs. deflection. Since it is not practical to computed as c = (q + s - l)/qs, where q = P/P,, conduct the test up to the large deflections s = P/P,,, P is the collapse load, P, is the local needed to realize the asymptotic value of load, buckling load (predicted or previously meaa data reduction technique (Tomblin, 1991) is sured on short columns) and P,, is the Euler buckling load (predicted or previously meaused. Columns having slenderness larger than 1.5 sured.) will buckle in a global mode (Euler).The Euler buckling load is controlled by the bending 46.3.3 REINFORCEMENT OF CONCRETE stiffness D. A column with pinned-pinned end conditions is subjected to an axial load (prefer- Reinforcement of Portland cement concrete ably under axial displacement control).Pinned can be accomplished with composite reinforcend conditions are the only conditions that can ing bars (rebars) instead of steel bars to be achieved with any degree of certainty. A minimize the corrosion of steel and cracking of 100% degree of fixity required by a clamped concrete caused by the expansion of the corend condition is not achievable on composite roding steel rebars. Composite rebars should columns because of the difficulties associated have good bond with concrete and adequate with connecting composites. Data from corrosion resistance. Pultruded rebars are the pinned-pinned columns can be used for other most common and least expensive alternative end conditions by using the effective length for the reinforcement (not pre-stressed) of conconcept (46.1). Weak axis tests are simpler to crete. Pultruded rebars have an angle perform, since strong axis tests require the use overwrap and/or a sand coating to improve of lateral support. Because of the imperfec- the bond with concrete. Fiberglass rebars, tions in the material and loading fixture, the aramid rebars, seven-wire carbon cables, load deflection plot has an hyperbolic shape, grids, and even gratings have been used in the buckling load being given by the asymp- research studies as reinforcements of concrete. Fiberglass rebars have lower modulus of tote of the plot. A data reduction technique elasticity than steel rebars, causing larger known as the Southwell method is used to
1000 Construction Creep is observed as increasing elongation under constant load. Relaxation is the reduction of stress over time for a constant elongation. Permanent levels of stress in the composite should be under the stress rupture limit (also called static fatigue) of the composite. Martine (1993) reported stress rupture of E-glass reinforced composite at 10 000 h at 58% of the initial strength. Glaser, Moore and Chiao (1983) report survival of S-glass reinforced epoxy specimen to a 10 year sustained load test: 90% of the specimens survived at 50% of static strength; 98% survived a 40% loading; and 100% survived a 35% loading. To account for stress rupture and other factors in the design of fully overwrapped pressure vessels for compressed natural gas, stress ratios (burst over service fiber stress) of 3.5 for fiberglass, 3 for aramid, and 2.25 for carbon reinforced composite were proposed by the American National Standards Institute (AGA NGV2 1992).These are indications that the static strength of fiber reinforcements cannot be utilized for long duration loads like those encountered in construction applications. Glass fibers, if unprotected, deteriorate when exposed to an alkaline environment. Therefore, glass reinforced composites rely on the protection provided by the resin to resist the attack of alkaline environments. Alkaline degradation is of particular concern in concrete reinforced with fiberglass composites since the concrete mix is alkaline. While concrete alkalinity is well documented at the time concrete is poured, only now are studies being conducted to investigate the level of alkalinity of concrete with time. Water is needed to establish a link between the glass fibers and the alkaline concrete environment. If water is present, the link can be established through cracks in the resin. Permeability of the poly46.3.4 ENVIRONMENTAL EFFECTS mer matrix resulting from voids or moisture Dead loads (e.g. the weight of the structure) diffusivity of the polymer are being considwill be applied to the composite material for the ered as possible additional mechanisms that life of the structure. Sustained loads induce two may place the alkaline concrete environment major effects in composites, creep and stress in contact with the glass fibers. rupture. Creep and relaxation are two altemaS2-glass reinforced Shell Epon 9310 comtive descriptions of the same phenomenon. posite rods produced by pultrusion were used
deflections than in steel reinforced concrete beams. Typically, the deflection of reinforced concrete (RC) beams is controlled by two factors: the amount of concrete that remains un-cracked (essentially in compression); and the modulus of elasticity of the reinforcement. The extension of cracking in the tension side of RC beams depends largely on the strength and uniformity of the bond between the rebar and concrete. Good bond translates into uniformly spaced cracks in the tension side of the beams. Therefore, a larger area of concrete remains un-cracked, thus contributing to the bending stiffness and limiting deflections. Fiber reinforced composite rebars have a linear stress-strain behavior up to failure, thus they experience a brittle behavior when compared to steel rebars that have a yield plateau (i.e. plasticity) before they rupture. Brittle behavior is not convenient from a safety point of view because of possible catastrophic RC member failure. However, concrete reinforced with fiberglass rebars experiences very large deflections up to failure, larger than comparable steel reinforced beams in their plastic regime, because of a combination of low modulus of elasticity and high tensile strength of the fiberglass rebars. No code or standard regulating the design of RC beams, reinforced with fiberglass composite reinforcing bars (rebars) is in place at the time of this publication. Faza and GangaRao (1993) suggest that it is possible to proceed along the lines of the American Concrete Institute guidelines ACI 318-89 complemented by the recommendations ACI 363 R, properly modified to account for the properties of the fiberglass rebars.
References 1001 by Sen, Issa, and Mariscal (1992) to pre-stress concrete piles. Seven 3 mm (0.125in) diameter rods were twisted with one turn every 30 cm (12 in) to create a seven wire strand. The strands were used to pre-stress concrete piles to be used in marine environment. To simulate marine environment, the piles were subjected to wet and dry seven-day cycles in a 15% sodium chloride solution, then tested to failure in bending. To facilitate the moisture ingress, four out of eight glass composite reinforced piles were pre-cracked at the mid-span prior to testing. The bending strength of specimens subjected to the wetting cycles, specially those pre-cracked, reduced significantly over time. From the pre-stressing force applied, it is predicted that 1.6% strain was induced in the rod, while larger values of strain may have been induced during the precracking of the beams. Additionally, radial stresses produced by the twist of the rods in the strand and by the moisture ingress in the resin may have caused cracking. Although the neat resin failure strain is reported at 4%, the strain applied to the composite rods may have produced cracking of the matrix, thus facilitating the ingress of moisture carrying the alkaline solution to the glass fibers. Scanning electron microscopy micro-photographs clearly show degradation of the glass fibers along with some cracking of the matrix. An additional pile was fabricated with pultruded fiberglass rods for which an additional coating resin was added to provide a smooth finish. This pile performed significantly better than the other four after a nine month exposure, perhaps because of the added protection of the finishing resin. It is concluded that the alkalinity of the environment (eg. concrete), the availability of water or other solvent, and the protection provided by the resin must be evaluated for each specific application. 46.4 CONCLUSIONS
Composite materials have been used for many structures that have performed well over
many years of service, mostly in adverse conditions. Composites are being used extensively for applications where the advantages significantly justify their higher initial cost. Composites have been successful when the design and manufacturing of a product was performed by a single company or group of companies, using the integrated design approach typical of the aerospace industry. The use of composites in the traditional civil engineering environment, where individuals create unique structures from standard components, has been slow, with the exception of markets where corrosion resistance offsets higher costs and less than optimum performance. Large markets have not developed partly because of lack of design codes and specifications and lack of awareness of composites advantages by structural engineers. A number of factors, including the decline of military industry and the need for rehabilitation of USA infrastructure, have produced a flurry of activity in this area, which undoubtedly will produce significant progress in the application of composites in construction.
REFERENCES Adams, R.C. and Bogner, B.R. 1993. Long-term use of isopolyesters in corrosion resistance. Proc. 48th Annual Conf, Section 1-C. New York: Composites Institute, The Society of the Plastics Industry Inc., pp. 1-5. AGA NGV2. 1992. Proposed American National Standard, basic requirements for compressed natural gas vehicle (NGV) fuel containers. Draft 8. American National Standards Institute. Ahmed, S.H. and Plecnik, J.M. 1989. Transfer of composite technology to design and construction of bridges. Washington, DC: US Department of Transportation, Federal Highway Administration, Office of University Research, Contract DTRS 5683400043. Bank, L.C. 1989. Flexural and shear modulii of fullsection fiber reinforced plastic (FRP) pultruded beams. J. Testing Evaluation, 17(1),4045. Barbero, E.J. and GangaRao, H.V.S. 1991. SAMPE J. Part I(12) 1991. Part II(1) 1992 Barbero, E.J. 1998. Introduction to Composite Materials
1002 construction Design. Washington, DC: Taylor and Francis. Barbero E.J. and Raftoyiannis, I. 1993. Local buckling of FRP beams and columns. ASCE J. Mater. Civil Engng, 5(3), 339-355. Barbero, E.J. and Tomblin, J. 1992. A phenomenological design equation for FRP columns with interaction between local and global buckling. Thin-Walled Structures. Composite Technol., 18, 117-131. Composite Technology Inc. 1992. Creative Pultrusions. 1989. Creative Pultrusions Design Guide. Alumn Bank, PA: Creative Pultrusions Inc. Davalos, J.F. and Salim, H A. 1992. Design of stresslaminated T-system timber bridges. Timber Bridge Information Resource Center, USDA Forest Service, Northeastern Area. Morgantown, WV. Faza, S and GangaRao, H.V.S. 1993. Pultruded fiber reinforced plastic bars, an alternative to steel reinforcement of concrete, Section 13-D, Proc. 48th A n n . Conf, pp. 1C-5C. New York: Composites Institute, The Society of the Plastics Industry, Inc. Fibergrate. 1992. Structural Fiberglass Products and Systems. Dallas, TX: Fibergrate. GangaRao, H.V.S. and Barbero, E.J. 1991. Construction: structural applications, in Encyclopedia of Composites (6), (ed. S. Lee). New York VCH Publishers. Glaser, R.E., Moore, R. and Chiao, T.T. 1983. Life estimation of an E-glass epoxy composite under sustained tensile loading. Composites Technology Review, 5(1), 21-26. Head, P. 1988. Use of fiber reinforced plastics in bridge structures. Helsinki: XI11 International Association for Bridge Structures Engineering (IABSE),pp. 123-128. Iyer, S.L. 1993. First composite cable pre-stressed bridge in the USA. Proc. 38th Int. SAMPE Symp. May 10-13, pp. 1766-1771. Iyer, S.L. (ed). 1991. Advanced Composite Materials in Civil Engineering Structures, (ed. S.L. Iyer). New York American Society of Civil Engineers. Kakihara, R., Kamiyoshi, M., Kumagai, S. and Noritake, K. 1991. A new aramid rod for the reinforcement of pre-stressedconcrete structures, in Advanced Composite Materials in Civil Engineering Structures, (ed. S.L. Iyer). New York American Society of Civil Engineers, pp. 132-142. Koga, M., Okano, M., Kawamoto, Y., Sakai, H. and Yagi, K. 1992. Application of a tendon made of
CFRP rods to a post-tensioned pre-stressed concrete bridge, in Advanced Composite Materials in Bridges and Structures, (eds. K.W. Neale and P. Labossiere).Montreal: The Canadian Society for Civil Engineering, pp. 405-414. Martine, E.A. 1993. Long-term tensile creep and stress rupture evaluation of unidirectional fiberglass-reinforced composites. Proc. Composites Institute 48th Ann. Conf., pp. 9-A/1-4. Meisseler, H-J and Preis, L. 1989. High performance glass fiber composite bars as reinforcements, in Concrete and Foundation Structures. Strabag BauAG information brochure. MMFG. 1992. EXTREN Fiberglass Structural Shapes Design Manual. Bristol, VA: Morrison Molded Fiberglass Co. Mochida, S., Tanaka, T. and Yagi, K. 1992. The Development And Application Of A Ground Anchor Using New Materials. Advanced Composite Materials in Bridges and Structures, (eds. K.W. Neale and P. Labossiere), Montreal: The Canadian Society For Civil Engineering, pp. 393-402. Mufti, A.A., Erki, M-A. and Jaeger, L.G. (eds). 1991. Advanced Composite Materials With Application To Bridges. Montreal: The Canadian Society for Civil Engineering. Mufti, A.A., Erki, M-A. and Jaeger, L.G. 1992. Advanced Composite Materials In Bridges And Structures In Japan Montreal: The Canadian Society for Civil Engineering. Munipalle, U.M. 1992. Analysis And Testing Of Wood-Glass Fiber Reinforced Plastic Adhesive Interface. M. S. Thesis. West Virginia University, Morgantown, W. Neale, K.W. and Labossiere, P. (eds). 1992.Advanced Composite Materials In Bridges And Structures. Montreal: The Canadian Society For Civil Engineering. Noritake, K., Kumagai, S., Mizutani, J. and Mukae, K. 1992. Construction of a pre-stressed barge using aramid FRP rods, in Advanced Composite Materials in Bridges and Structures, (eds. K.W. Neale and P. Labossiere), Montreal: The Canadian Society For Civil Engineering, pp. 533-541. Pletcher, D. 1991. Consider structural composites, Chem. Eng. Progress, November, 4449. Porter, M.L. and Barnes, B.A. 1991. Tensile testing of glass fiber composite rod, in Advanced Composite Materials In Civil Engineering Structures, (ed. S.L. Iyer). New York American Society Of Civil Engineers, pp. 123-131.
References 1003 Priestley, M.J.N., Seible, F. and Fyfe, E. 1992. Column seismic retrofit using fiberglass-epoxy jackets, in Advanced Composite Materials in Bridges and Structures, (eds. K.W. Neale And P. Labossiere). Montreal: The Canadian Society For Civil Engineering, pp. 287-298. Raftoyiannis, L. 1993. Buckling Mode Interaction in Fiber Reinforced Composite Structures. Ph.D. Thesis. West Virginia University, Morgantown, WV. Rostasy, F.S., Hankers, C. and Ranisch, E-H. 1992. Strengthening of R/C- and P/C-structures with bonded FRP plates, in Advanced Composite Materials in Bridges and Structures, (eds. K.W. Neale and P. Labossiere). Montreal: The Canadian Society For Civil Engineering, pp. 253-263. Saadatmanesh and Ehsani. 1990. Fiber composite plates can strengthen beams. Concrete International. Farmington Hills, MI: American Concrete Institute, pp. 65-71. Sen, R., Issa, M. and Mariscal, D. 1992. Feasibility of Fiberglass Pretensioned Piles in a Marine Environment. Report No. CEM/ST/92/1, University of South Florida, Tampa, FL
Gere, J.M. and Timoshenko, S.P. 1990. Mechanics of Materials, 3rd Edn. Boston: PWS-KENT Publishing Co. Tomblin, J. 1991. A Universal Design Equation For Pultruded Composite Columns. M.S. Thesis. West Virginia University, Morgantown, WV. Tomblin, J. 1994. Compressive Strength Models For Pultruded Fiber Reinforced Composites. Ph.D. Dissertation, West Virginia University, Morgantown, WV. TUFSPAN. 1991. Technical Data and Design Guide, Forth Worth Tufspan, p. 8. Wolf, R. and Miesseler, H-J. 1992. Experience with glass fiber pre-stressing elements for concrete bridges, in Advanced Composite Materials in Bridges and Structures, (eds. K.W. Neale And P. Labossiere). Montreal: The Canadian Society For Civil Engineering, pp. 425433. ACKNOWLEDGEMENTS
My sincere gratitude to Prof. G. Turvey for his help researching applications in Europe, to all the contributors of information for this article, and to West Virginia University for the support of this project.
AEROSPACE EQUIPMENT AND INSTRUMENT STRUCTURE
47
G a y C. Krumweide and E d d y A. Derby
47.1 INTRODUCTION
aluminum, or very expensive beryllium further justified the material selection. It is necessary to discuss GFRP (graphite fiber Consequently, numerous mirror bezels, reinforced plastics) thoroughly in order to telescopes, optical benches and reflectors were understand the importance of composites in designed and built from GFRP in the early the manufacture of aerospace equipment and 1970s. As a result of these efforts, today more instrument structures. Before composite mateand more structures are being fabricated from rials such as GFRP became viable candidate GFRP materials, principally because specific materials for aerospace primary structures, or even for sporting equipment, they were first materials matching specific property requireused for aerospace equipment and instrument ments are now available. Although GFRP has dominated composite structures. Because composites exhibited both materials applications, DuPont's Kevlar-49 unique and superior properties, designers has been found to be ideal for antenna reflecwere willing to pay the prevailing high prices tors because of its extremely light weight and to achieve their design goals. For space hardRF transparency. Many communication satelware, where a pound of weight saved was lites utilize this type of Kevlar reflector, such worth thousands of dollars, designers were as the SatCom-F, Telstar, ANIK-E, SpaceNET, motivated to characterize composite materials and Superbird SCS (Fig. 47.1). suitable to their applications. Understandably, composites for primary and secondary structure (e.g. launch vehicles, aircraft frames, wing spars and skins, etc.) were a 'hard sell', and temperature extremes were too severe for 'thermosetting' GFRP to be used on missiles. The quantity of GFRP required for a particular piece of aerospace equipment or instrument structure was usually minimal, so relative costs were low, and composite materials' superior properties compared to heavy Invar, or high coefficient of thermal expansion (CTE)
Handbook of Composites.Edited by S.T. Peters. Published in 1998 by Chapman & Hall,London. ISBN 0 412 54020 7
Fig. 47.1 SuDerbird SCS Kevlar dual-shell reflector.
Historical perspective and progress 1005 For aerospace equipment and instrument structures, generally, thermoplastic versions of GFRP have not seen as much application as the thermosets due to high investment costs in tooling and facilities. Considering the relatively small quantities that are usually bought, thermoplastic applications are not often cost effective. Also, the required high temperature cures subject the laminate to microcracking instabilities. Metal matrix composite applications have been limited, and knowledge of material properties and processes has been restricted due to their use on classified programs. Materials like silicon carbides and carbon-carbon find limited, but important, use in aerospace structures, particularly mirrors. Design selection criteria plays an extremely important role in determining what type of material is used on any particular component of aerospace equipment and instrument structures. The primary reasons for choosing a composite material, rather than a metal, are weight, dynamic stability and thermal stability. Table 47.1 categorizes some of the desirable and undesirable characteristics that confronted the early users of GFRP. Table 47.2 addresses the undesirable properties listed in Table 47.1 and indicates how the aerospace equipment and instrument designers have recognized the barriers or problems with GFRP and found work-around techniques to allow their usage. Table 47.3 illustrates several structural applications where GFRP have been used
and/or are going to be used for aerospace equipment and instrument structures. Table 47.4 compares the mechanical and thermal properties of candidate materials for aerospace equipment and instrument structures. One could ask, from the obvious property advantage of beryllium, why all such structures are not made from beryllium? If not beryllium, why not metal matrix composites or carbon-carbon (C/C) or silicon carbide (Sic)? The answer is that raw material cost, fabrication cost, practical size, and other critical properties all come into play for any specific application. Table 47.5 shows typical design requirements for various applications and indicates which materials typically satisfy the critical requirement. The primary reason that GFRP is a material candidate for most applications is the wide range of material systems that are currently available that can compare with Kevlar, aluminum, Invar, beryllium, metal matrix composites, silicon carbide, and carbon-carbon. 47.2 HISTORICAL PERSPECTIVE AND PROGRESS
From a historical perspective, and to illustrate where significant progress has been made in the use of composites for aerospace structure, the following areas of interest will be addressed:
Table 47.1 GFRP properties Desirable
Low density Coefficient of thermal expansion(CTE near zero) High specific strength High specific stiffness Readily formable Crack growth resistant Adaptable laminate properties Easily repairable
Undesirable
Low short transverse properties Hygroscopic High material cost High fabrication cost Low impact strength Subject to microcracks (translaminarstress relief) Low peel strength
1006 Aerospace equipment and instrument structure Table 47.2 GFRP design barriers and work-around techniques
Undesirable properties
Barrier to design
Work-around techniques
Reference
Anisotropic behavior
Low strength in angles
1. Radius blocks 2. Substitute material (metals)locally 3. Use mortise and tenon joints 1. Use mortise and tenon joints 2. Butt edges/insert 3. Back-to-back splice 1. Egg crate joint 2. Mortise and tenon 3. Add local compensation 4. Maximize in-plane material orientation 5. Fitting Boss through
Stumm,1981 Campbell, 1981 Krumweide, 1988
Thermal instability of angles High throughthickness expansion of laminate
Stumm, 1981 Campbell, 1981 Krumweide, 1988 Stumm, 1981 Krumweide, 1988 Krumweide, 1991
Skin
Hertz, 1977; Stumm, 1979 Walrath, 1979; Levy, 1984 Krumweide, 1989 Telkamp, 1990 Krumweide, 1991; Brand, 1992
Hygroscopic nature
Expansion/distortion
1. Moisture barrier 2. Define exposure/ drying scenario 3. Cyanate ester
Material cost/ fabrication cost
High basic part cost
Susceptibility to microcracks
Micro properties affected
Susceptibility to impact damage
Low impact strength
Low peel strength
Low joint allowables
1. Proper material Krumweide, 1977 selection Krumweide, 1988 2. Tooling/fabrication techniques 3. Assess cost/weight ratio 4. Ovencure 5 . Minimize pieces 6. Eliminate molded parts 7. Tailor joints/load Stumm, 1981 1. Uselow temperature Krumweide, 1991 curing resin systems Brand, 1992 2. Low stacking angle 3. Small ply thickness 4. Reduce fiber modulus 5 . Thermal cycle parts 6. Cyanate ester Dunbar, 1978 1. Protect surface with Kevlar/honeycomb Herrick, 1984 2. Material design (thickness, orientation) Dunbar, 1978 1. Use mortise and tenon joints Krumweide, 1979 Stumm, 1981 2. Fasteners and angle clips
Historical perspective and progress
0
0 0 0
problems encountered/work-around techniques; construction methods; tooling development; major milestones.
47.2.1 PROBLEMS ENCOUNTERED/WORKAROUND TECHNIQUES
Aerospace designers have constantly been seeking the ideal material - the one material which meets design requirements with the fewest compromises. Of course, few materials are completely ideal for a given set of requirements and design trade-offs are almost always necessary. Table 47.2, above, summarizes typical problems, presents some design concerns, and/or 'barriers to design' associated with compos-
1007
ites, principally GFRP, and indicates a variety of work-around techniques which have been implemented successfully to address these design concerns. To help explain and clarify these techniques, some examples of specific design solutions are discussed below. Moisture effects For GFRP, it has been determined that the predominant mechanism of water penetration into the laminate is through the resin by a diffusional process. The moisture diffusion is governed by Fick's second law which is similar to Fourier's equation for thermal conductivity:
6%
6C
sf
=
Dzsz'
(47.1)
Table 47.3 GFRP Applications for aerospace structures
Equipment structures
RF Systems: -Reflectors (Viking, Nimbus-G, ACTS) -Feed horns and waveguides (INTELSAT) -MUX cavities -Diplexers -Phased arrays
Instrument structures Metering structures (LANDSAT, Thematic mapper; SOHO, UVCS; HEAO-B, COSTAR and Hubble Telescopes) Camera housings (Mars Observer Camera (MOC), Hubble WFOV Camera)
Large segmented reflectors
Optical benches (MAGSAT, UARS HRDI, Hubble FGS Optical Benches)
Solar panel (substrates) (BS-3, RADARSAT, Mars Observer)
FPA or relay optics (Hubble Telescope; THEMATIC MAPPER)
Booms, Stardust (Shuttle RMAB, M-SAT)
Mirror bezel
Bus (Mighty Sat M I , SMEX/WIRE, Indostar Forte)
Mirrors (Microwave Limb Sounder, NGST, INM Scan Mirror)
Electronic chassis (EO1, SMTS-OBC, MARS 98, Stardust)
Lens holders
CCA cardguides (NAWC Coldplate)
Support benches (Hubble FGS)
Solar Concentrators
Startrackers (Hubble Equipment Shelf) Submillimeter reflector Helicopter mast mounts Laser comm gimbals
1008 Aerospace equipment and instrument structure where c = moisture concentration; t = time; Dz = moisture diffusivity; and z = thickness coordinate. Diffusivity is dependent on resin type and temperature but is independent of moisture concentration or laminate orientation. Figure 47.2 illustrates the importance of temperature in determining the diffusivity of a laminate. Figure 47.3 illustrates the moisture absorption behavior as a function of temperature for this same laminate. Note that this is for a 100% relative humidity (RH) exposure. Typical moisture absorption values (primarily by the resin) are 3 . 5 4 % by weight for epoxy and 1-1.3% by weight for cyanate ester. The
maximum moisture pick-up for the composite (Mcm)may be calculated by the following equation:
MCm= (Mrn)[x
wr
(47.2)
where Wr = resin weight YOand (MJr is the maximum moisture pick-up for the neat resin for a given relative humidity condition. The moisture content as a function, Iw,can be represented by: Mcm= A
(WB
(47.3)
where A and B are coefficients obtained from empirical data. Figure 47.4 depicts the percent of moisture pick-up for various levels of
Table 47.4 Typical mechanical and thermal properties of isotropic/quasi-isotropic materials ~~
~
Material
Copper Beryllium Aluminum Titanium Stainless steel Super Invar Fused silica Ule fused silica Zerodur P75S/EP P100S/EP P120S/EP PlSOS/EP T300/EP T50/EP AS-4/EP XN50A/EP XN70A/EP M55J/EP M60J/EP FT500/EP FT/ 700/EP GY70/EP Boron/EP S-GL/EP KV49/EP KV-149 NA = Not Available
Modulus E (GPa) 117 289 70.3 113.7 193 144.7 73 67.5 90.3 103.3 143.3 170.9 186 55.1 82.7 51.7 103.4 144.7 103.4 120.6 82.7 137.8
Density (Mg/m3) 8.86 1.83 2.68 4.43 8.03 8.03 2.19 2.19 2.54 1.72 1.80 1.83 1.83 1.58 1.58 1.58 1.80 1.83 1.66 1.69 1.77 1.80
110.2
1.66
89.6 26.9 29.6 40.6
2.05 2.02 1.38 1.38
CTE
CP
( X 1O-6/K)
(kJhgK)
16.6 11.5 23.8 9 16.5 0.18 0.5 0.029 0.108 -0.14 -0.72 -0.85 -1.08 2.7 0.72 2.52 -0.072 -0.54 -0.27 -0.45 0.36 -0.72 -0.18 5.4 9.7 6.3 1.8
0.38 1.88 0.96 0.54 0.50 0.50 0.71 0.75 0.92 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.71 NA NA
K (WImK) 398 179-207 138-237 16.9-20.8 16.26 13.8 1.33 0.86 1.7 55.3 116 190 334 2.4 20.8 5.2 48.4 98.6 20.8 24.2 41.5 81.3 NA NA 0.34 1.04 1.04
Historical perspective and progress 1009 Table 47.5 Material selection criteria
Aerospace application Critical requirements
-
Equipment structure
Instrument structure
RF reflectors
Solar panel substrates
Bus structures
Booms
Stable structures
Mirrors
Mass
GFRP (1) Kevlar
GFRP(1) Kevlar
GFRP Aluminum M/M
GFRP Aluminum Beryllium M/M
GFRP (5) Beryllium
GFRP (6) Beryllium Si/C, C/C
Dynamic stability
GFRP (1) Kevlar
GFRP (1) Kevlar Aluminum
GFRP Aluminum M/M
GFRP Aluminum Beryllium M/M
GFRP (5) Beryllium Invar
GFRP (6) Beryllium Si/C, C/C Aluminum
Thermal stability
GFRP
GFRP (1) Kevlar
GFRP (5) Invar
GFRP Beryllium M/M
GFRP (5) Invar Beryllium
GFRP (6) Invar Beryllium Si/C, C/C
Dynamic loads (Gs)
GFRP (1) Kevlar
GFRP (1) Kevlar Aluminum
GFRP Aluminum M/M
GFRP Beryllium M/M Aluminum
GFRP (5) Beryllium Invar
GFRP (6) Invar Beryllium Aluminum
Temperature extremes
GFRP (2) Kevlar
GFRP (2) Kevlar Aluminum
GFRP (2) Aluminum M/M
GFRP (2) Beryllium M/M Aluminum
GFRP (2) Invar Beryllium
GFRP (2x6) Si/C, C/C Beryllium Aluminum
Hygrostability
GFRP (2)
GFRP (2) Aluminum
GFRP (2) Aluminum M/M
GFRP (2) Beryllium M/M Aluminum
GFRP (2) Beryllium Invar
GFRP (2)(6) Invar Beryllium Aluminum Si/C, C/C
High thermal conductivity
GFRP (3)
GFRP (3) Aluminum
GFRP (3) Aluminum M/M
GFRP (3) Beryllium M/M Aluminum
GFRP (3) Beryllium Invar
GFRP (3x6) Invar Beryllium Aluminum Si/C, C / C )
Low thermal conductivity
GFRP (4)
GFRP (4) Kevlar
GFRP (4)
GFRP (4)
GFRP (4)
GFRP (4x6)
RF transmissibility Kevlar
Kevlar
-
-
-
-
cost
GFRP (1) Kevlar Aluminum
GFRP Aluminum
GFRP Aluminum
GFRP Invar
Aluminum Invar
GFRP Kevlar
~
(1)Combinations of Kevlar and GFRP is used when mass and dynamic stability is important. (2) Cyanate resins have been shown to handle temperature extremes (T, high, no microcrackhg). (3) Pitch fibers (especially ultra, ultra high modulus have high thermal conductivity).
(4)PAN fibers have low thermal conductivity. (5)Metals shown may be applicable if a small size structure. (6) GFRP has been used successfully as mirror substratesand for some submillimeter reflectors core and skin.
1010 Aerospace equipment and instrument structure or
A L / L = PAM E
I.WOE4
'-
4- j I i j I l j * l ~
0.00211 0.0032 O.WJ6 TEMPERATURE( 1 I K )
0.0024
0 . W
Fig. 47.2 Diffusivity against temperature for P75S/cyanate; P75S/epoxy system.
relative humidity and shows the significance between an epoxy and cyanate ester resin. Table 47.6 provides values for the swelling coefficient @) for typical GFRP materials used to fabricate aerospace structures, where p is calculated from:
p
=
(E&)
Once /3 is known, the expected strain can be calculated for a given percent moisture content in the laminate. The significance of Fig. 47.4 cannot be overemphasized. Up until 1989 designers of aerospace structures had to utilize the work-around techniques 1and 2 mentioned in Table 47.2 in order to handle the excessive expansion and the corresponding distortion associated with the hygroscopic nature of epoxy based GFRP material systems. Table 47.7 indicates work-around techniques utilized on various programs before cyanates were introduced. Cyanate ester resin systems now being used (Brand et al., 1992) in GFRP materials offer increased stability at reduced cost for aerospace structures, primarily due to the possible elimination of expensive moisture barriers (10% of manufacturing costs) and the even greater cost of facilities for dryout processing.
(47.4) IW
-
0 40
rm
ERLWW RES4N (
E
g
0 30
worn I
P76S I ERLISBZ
080
(EwxIl
w
5
080
P
I s
URE
z
'
P765 I ERLlOsO
J
I CYUUTE I
L
020
0 10
OW 0 00 0 00
2 00
4 00
8 00
8 00
1000
RELATIVE HUMIDITY I X )
TIME ( H O U R S ~ 7 I P )
Fig. 47.3 Percentage moisture change against time (h1Iz)for 0 . 5 m thick P75S/cyanate at 100% RH exposure.
Fig. 47.4 Percentage moisture pick-up against relative humidity (Yo). (a) ERL1999 epoxy resin; (b) P75S/ERL1962 epoxy; (c) P75/ERL1999 cyanate.
Historical perspective and progress 1011 Table 47.6 Typical material CMEs for pseudoisotropic laminates
p (PPM/%M)
Material T300/EP T50/EP P75S/EP P100s / EP P120S/EP P75S/CY
400 215 162 103 86 105
M , @ 100% RH 1.13 1.13 1.03 1.03 1.03 0.32
Figure 47.5 illustrates the effect of composite microcracking. As the temperature is lowered Microcracking of a composite which causes and reaches the threshold of microcracking, as dimensional instability should no longer be an evidenced by an abrupt strain change, the issue if the proper material selections are curve changes to a new slope. Dimensional made. A need to thermal cycle to gain dimeninstability manifests itself in the following sional stability (causing translaminar cracking ways: or microcracking to progress to completion) was a requirement in the past, e.g. in Hubble 0 A hysteresis effect is produced in the structure. Telescope tube components. This was due, primarily, to of the relative brittleness 0 Changes in CTE occur (CTE becomes more negative with increased thermal cycling). (non-toughened) of the resins and the unavailability of thin cure ply thickness prepregs. 0 Moisture response rates and moisture levels increase. Also, at the time, the phenomenon was not 0 Irreversible expansion of the GFRP material. fully understood.
Microcracking
Table 47.7 Prior work-around technique for addressing the hygroscopic nature of graphite/epoxy Annotations (Ref.)
Instrument
Work-around technique (from Table 47.2 Hygroscopic nature)
Teal ruby
1
Stumm, 1979. Cryogenic application, indium/bismuth deposited metal-type moisture barrier - space application
Thematic Mapper
2
Walrath, 1979. Two flight units operational, bake-out and environmental controls during assembly and test very well controlled and monitored - space application
Mars Observer camera
2
Telkamp, 1990. Pre-flight unit test provides dimensional change data, compensated for by preset of optics - space application
HRMA Cylinders HEAO-B
1
Hertz, 1977. Aluminum foil-type moisture barrier; space application
1012 Aerospace equipment and instrument structure I
2%
Anisotropic behavior The anisotropic behavior of GFRP offers a challenge in the design of aerospace structures. Some typical areas of concern are joints, springback, cutouts in cylinders and bowing of panels and assemblies.
Joints Early in the 1970s, designers discovered that the through-the-thickness (translaminar) properties exhibited by GFRP reduced strength and increased CTE values for joints, Fig. 47.5 Interpretation of common events occur- i.e. the CTE in-plane may be 0.10X10-6/oCand ring during measurements illustrating 34X104/"C through the thickness. In the folmicrocracking of pseudoisotropic GY-70/930 lowing years, however, improved (0"/45" /90° / 135")%. 12mm/ply. mechanical/ thermal joint designs have evolved which optimize joint characteristics relative to thermal distortion/ cost/weight requirements. Figure 47.6 illustrates this evolution for both fixed and removable joints with the development path moving from A to D. Today the following procedures can be used by designers to select appropriate fiber and resin systems to eliminate or minimize microcracking. 0
0
0
0
0
Establish the thermal range of the exposure environment. Determine the stiffness, CTE, and weight requirements of the application. Use a thinner cure ply thickness (CPT) prepreg, if high modulus materials are required (e.g. 0.0063 mm or less CPT) or a woven fabric. Select a resin system compatible with the thermal and mechanical requirement (e.g. cyanate esters such as Fiberite 954-3 or YLA RS-3 do not microcrack when used with UHM pitch fibers (Amoco P75S) even at cryogenic temperatures). Establish the optimal fiber orientation (e.g.
@ WAL
OCUBLE CLIP
@ MllY
(0"/45°/900/1350)s).
Of course, adequate testing of the material laminates before they are employed in the structure will ensure that correct material selections have been made.
Fig. 47.6 GFRP joint design evolution.
6 TENON
Historical perspective and progress 1013 Springback
Flat panel bowing/warping
Springback occurs when composite laminates are formed into angles during processing. Springback, or angle closure, occurs to the extent of about 1.5" upon removal of the laminate from the tool (General Dynamics Corporation, 1985). Exposure to moisture conditions will open this angle (via moisture absorption) to return to, or even exceed, the original angle. Through-the-thickness expansion of the laminate is the source of this dimensional instability. Using flat laminates and butt bonding at right angles is the workaround technique. Very thin clips can be used at the joint. Thermally, structures can be very sensitive to this problem as was discovered in the late 1970s in the testing of the GEMS Forward Mirror Support (Campbell et al., 1981). This problem has often been overlooked by designers over the years and needs to be considered whenever composites are utilized on dimensionally stable structures. Back-to-back angles or more simply mortise/tenon joints have been the work-around technique.
When laying up a flat panel, the typical fiber angle tolerance from the desired angle is O G o . This manufacturing tolerance is a major contributor to the bowing or warping of a panel. The thicker the laminate, the more a problem these variations are, in that substantial forces are required to make the panel flat. The laminate can straighten upon moisture absorption, and drying the laminate to its original cured state will bring back the prior laminate geometry. The best solution for e h i n a t i n g bowed or warped panels is to use a 'rotate and fold' laminate technique, specifically developed, to produce essentially balanced laminates. This technique is depicted in Fig. 47.7. In essence, ply lay-up tolerances are canceled by becoming symmetrical, with the resulting laminate properties truly isotropic throughout the laminate. Note that this process is only applicable to pseudoisotropic laminate configurations.
Cut-outs in curved surfaces (cylinders) Some design approaches require holes in cylinders. The degree of distortion when a large hole is cut in a cylinder is surprisingly extensive and the stresses generated can force the cylinder to take a different shape (oval or hour-glass). Even the ends of a cylinder that has a large hole in it will not be flat or parallel to each other. The thicker the cylinder wall, the more difficult it is to correct this condition. Stiffening rings usually correct the problem on thin walled cylinders. Springback and cutouts, in formed parts, may reduce the benefit gained by a reduced number of piece parts if tight dimensional control is necessary. Parts may not fit together. Again, through-the-thickness expansion of the laminate is the source of this dimensional instability.
Fig. 47.7 The rotate and fold laminate construction technique.
1014 Aerospace equipment and instrument structure
Assembly bowing By maximizing the use of flat stock and making dual skin, eggcrate core structures, very stable assemblies are possible. This brings about an additional concern with through-thethickness expansion in that when an egg-crate core is used in a structure, severe bending is to be expected unless a specific design technique is employed. Since through-the-thickness expansion is 1000-times the basic in-plane expansion of the laminate, bending of the structural assembly occurs. Figure 47.8 illustrates the correct and incorrect slotting method. If done correctly the structure (Fig. 47.9) stays relatively straight through thermal Fig. 47.9 FGS keel. The structure remains stable cycling and moisture absorption. because the correct slotting method was employed. Mortise and tenon joints, used instead of clips, also promote stability. INCORRECT SLOTTING METHOD
I
1
n
n
n
WET OR T
n
1
I
R.T.
lnnnnl ORY OR T
=
47.2.2 CONSTRUCTION METHODS
Some important design options for aerospace equipment and instrument structure are: 0 0
R.T.
0
CORRECT SLOTTING METHOD
I
U
n
U
WET OR T
R.T.
DRY OR T
: R.T.
0
n
CROSSING RIB THICKNESS ADHESIVE
Fig. 47.8 Correct and incorrect slotting methods.
truss (tubes) or cylinders (for telescopes); monocoque skins or honeycomb sandwich (for reflectors); bonded, or bonded and bolted (for typical joints); molded unibody or flat laminate/bond assemblies (for typical structures).
Truss compared with cylinder (telescopes) Material availability, size, weight, interface, loads and stiffness requirements or considerations may drive the designer to a particular configuration for a particular application. Designers who favored tubular truss structures in the 1970s might not choose such an approach today. Design techniques are different, material choices are many, and manufacturing methods have changed. Today, for example, the Hubble Telescope could be designed as a faceted, dual-shell structure with a ribbed core. The faceted dual-shell (skins) could have water-jet cutouts with a
Hisforical perspective and progress latticework appearance. Then, again, the telescope might look the same but be fabricated from a non-microcracking, low moisture absorbing, thin-prepreg, cyanate ester prepreg system material.
Monocoque skins compared with honeycomb sandwich (reflectors) Large GFRP honeycomb reflectors were successfully fabricated in the mid-1980s for the Advanced Communication Technology Satellite (ACTS) spacecraft reflector. At the same time that large membrane reflectors were made for a Direct Broadcast Satellite (DBS) spacecraft reflector. They both met the desired design requirements. The tie-down configurations and weight constraints, in part, dictated which approach was utilized. Here again the requirements of size, weight, interface, loads, stiffness and material availability will influence the design. Designer experience and preference will also determine what concept is used in the future.
1015
a molded part is excessive on a particular design, a molded part may not be better. Then again, flat laminate construction may mean that more parts have to be made, inspected and handled, and if this cannot be done efficiently by the manufacturer, then it may not be the best approach. Some manufacturers can handle flat laminate construction very efficiently, if they have experience of assessment of cost, weight, risk. Many flat laminate 100% bonded assemblies for aerospace equipment and instrument structure have been fabricated and successfully flown, both on aircraft and spacecraft. Apparently the previous high risk, thought to exist with this approach, has not been proven true in practice. Here again, designer experience and preference will determine the course to follow. 47.2.3 TOOLING DEVELOPMENT
The number of cures, the size of the part and facilities available influence the choice of tooling to such a degree that each product fabricated must be thought-out thoroughly to Bonded joints compared with bonded and select the proper tooling. bolted joints An advantage of flat laminate construction Experience with all bonded structures, cost is that it eliminates a need for fabricating proconsiderations, and successful application (i.e. duction molds. A designer may want to look at hardware on-orbit) are driving designers of this approach first in order to cut costs, if expeaerospace structures away from fasteners alto- rience and fabrication techniques support this gether. The short cyclic load duration for method of fabrication. The high cost of production molds can be spacecraft today are such that little, if any, benefit is gained from bolts or fasteners and their significantly reduced through the use of alteruse should be minimized. Aircraft applica- nate techniques. For example, for cylinders, tions for equipment and instruments may be thin-walled rolled and welded aluminum another story and each application needs to be molds create dramatic savings as their light reviewed thoroughly, assessing cost, weight, weight permits envelope bagging which eliminates the need to withstand autoclave risk and so on. pressures. Monolithic (bulk graphite) molds are Molded (unibody) compared with flat expensive and heavy (storage and heat-up laminate bondedassemblies rate concerns), but their CTE match to GFRP The use of fewer parts may not be beneficial or Kevlar/epoxy is a great advantage. if they do not match up well at assembly. Greater accuracy and better replication of Also, if the cost of tooling and touch labor for parts is possible.
1016 Aerospace equipment and insfvument structure
The tooling concept and materials used to fabricate the tooling can be as important to the success of the hardware as the hardware design itself. The aerospace structural designer must stay in touch with technological developments in tooling and tooling materials. 47.2.4 MAJOR MILESTONES
Milestones in the use of composite materials have been in the areas of materials develop-
ment and their application to aerospace equipment and instrument structures (Tables 47.8 and 47.9). The photograph of the Hubble Space Telescope (Fig. 47.10) only shows the metering truss structure for the telescope; substantially more GRFP was used in the fabrication of the Ford Plane Assembly (FPA) and the many other components. The milestones cited represent major advances and point the way for continued use of composites for a wide variety of aerospace structures.
Table 47.8 Milestones for GFRP materials Material
Achievement
Thin prepregs
Reduced weight of hardware components Minimized microcracking Supported mirror technology Increased thermal stability of laminates
Toughened epoxies
Increase impact resistance of laminate Increase bond strength (interlaminear) of joints Increase compressions strength of laminate
High modulus PAN fibers
Reduce hardware weight Allow high stiffness, thermal stable, cyanate laminate Allow high strength application
Ultra-ultra high modulus pitch fibers
Increase thermal conductivity Minimum weight, stiff structures Increased EM1 capability
Cyanate ester resins
Reduced moisture levels Reduced strain response to moisture level Reduced microcracking Increased use temperature (high TE)
Table 47.9 Milestones for GFRP applications Application
Date
Achievement
Hubble telescope
1970-1980
Design and fabrication of a very large (2.4 m diameter x 5.2 m length) and thermally stable structure
Nimbus-G
1970-1980
No fasteners - membrane shell 100% bonded microwave reflector; Qual unit qual-tested twice. Flown in 1990 as TOPEX.
Shuttle RMAB
1970-1980
Reusable structure. Multiple shuttle missions.
Current applications 1017 -
Fig. 47.11 ACTS reflector. Fig. 47.10 Hubble space telescope metering truss assembly.
47.3.2 HRDI OPTICAL BENCH (UARS)
This 1.24 x 1.8 x 0.11 m thick, 49.9 kg optical bench for NASA's Upper Atmosphere Research Satellite (UARS) mission was nearly 47.3 CURRENT APPLICATIONS 100% bonded (Fig. 47.12). The only fasteners were anti-peel fasteners at the four comers 47.3.1 ACTS REFLECTORS where graphite/epoxy (T50/ERL1962) panel The Advanced Communications Technology sides terminate at titanium fittings. The bench Satellite (ACTS) reflectors were manufactured carried 3.2 times its own mass and had a CTE by Composite Optics, Inc. (COI), for General of 0.36 X 10-6/oC(Dodson, 1989). Electric/Astro-Space Division under contract with NASA Lewis Research Center (Fig. 47.11). 47.3.3 WCS(SOH0) These large reflectors (2.2 and 3.3 m aperture) had measured surface RMS accuracies of The Ultraviolet Coronagraph Spectrometer approximately 0.071 mm over their entire sur- (WCS) for the European Space Agency's Solar face after thermal cycling. These reflectors and Helospheric Observation (WHO) spacewere launched into orbit in September 1993 craft (1995 launch) is 100% bonded and are functioning well. Their GFRP skins graphite/epoxy (P75S/ERL 1962) (Fig. 47.13). (P75S/ERL1962) and Kevlar honeycomb core This truss structure is the first of its kind in that (Kevlar 49/934) resulted in a 3.44 kg/m2 the basic truss panels are bonded assemblies weight for the completed reflector (Rule, 1989). that have an I beam cross section. Eighty
1018 Aerospace equipment and instrument structure
ophcalstabilitynquktnmts(CTE-0.09 X lW/OC) had to be satisfied along with strength and frequency (70 Hz) requirements. The structure will hold 90.7 kg of equipment and only weighs 21.8 kg itself (Kilpatrick et al., 1990).Most of the surfaces on the structure were clearable and inspectable (minimal closed areas). 47.3.4 COSTAR (HUBBLE TELESCOPE CORRECTIVE OrrrCS)
Fig. 47.12 HRDI bench (UARS).
The Corrective Optics Space Telescope Axial Replacement (COSTAR) was developed by Ball Corporation by NASA/GSFC. This structure (Fig. 47.14) supports a Deployable Optical Bench (DOB)that serves as the corrective optic for the Hubble Space Telescope. Built by Hercules Aerospace Company, this composite structure has to have a low CTE (<0.15) and
:ig. 47.13 W Coronagraph Spectrometer (UVCS) telescope structure assembly (SOHO).
Predictions
1019
Fig. 47.14 Corrective Optics Space Telescope Axial Replacement (COSTAR).
high resonant frequency (100 Hz for the DOB). The long-term stability effects due to moisture absorption (ground) and desorption (space) were addressed by optimizing the composite laminate orientation. The long-term CTE concerns were more important in meeting line-of-sight requirements (>15 arc-sec) than was a need to optimize CTE. A removable cover and mounting of aluminum electronic box required very unique design solutions (Neam and Gerber, 1992). 47.4 PREDICTIONS
The dominance of GFRP as the preferred choice composite material in the fabrication of aerospace equipment and instrument structures will continue into the next century because of a number of reasons:
0
0 0 0
new materials; new manufacturing processes; new applications; economics.
This, of course, does not mean other composite materials will not be used, but their use will be in similar proportion as is the current practice. 47.4.1 NEW MATERIALS
GFRP is a constant evolution of material variations. Fiber makers continue to change their fiber manufacturing processes to create a new fiber to satisfy the needs of the user. Properties like fiber strength, modulus, and thermal/electrical conductivity are continuing to be adjusted for various application requirements. The demand for properties changes also are requested of the resin manufacturers and/or
1020 Aerospace equipment and instrument structure prepreg suppliers. Cyanate ester resin formula- 47.4.2 ECONOMICS tions are in their infancy, current GFRP A 1993 conference held in Logan, UT, sponcyanates have levels of moisture absorption sored by the AIAA and Utah State University, less than one-third the GFRP/epoxies. Soon had a basic theme for small satellite structures moisture absorption for GFRP/cyanate may be of ‘cheaper, faster, better’. The ’better’ was in one-tenth that of GFRP epoxies. In the future, some cases interpreted as lighter for better we may have thermoset resins that will have performance. no moisture absorption and their operating One of the primary messages to the industemperatures may be over 240°C (464°F). try from that conference was that economics dictates that satellites become smaller and lighter so that the ’on-orbit’ cost for a particuNew manufacturing processes lar satellite is much less, can be launched by a It is hard to imagine new manufacturing meth- small, low-cost launch vehicle (e.g. OSC ods for processing GFRP that are not in use Pegasus or Lockheed LLV1). GFRP is currently today, but the demand for less expensive the most logical choice to satisfy the requirematerials, and hardware may bring develop- ment of cheaper, faster, and better. ments. One method currently being developed Manufacturing processes are developed that by Composite Optics, Inc. in San Diego, CA, will allow lightweight GFRP to be utilized for called SNAPSAF is predicted to reduce the bus structures, solar panel substrates, reflectors manufacturing cost to less than half current and instruments. Weight reductions of 5040% levels. are possible with GFRP. Package size may be the only limitation to allowing three or four satellites to be launched rather than one or two. New applications Creative packaging and miniaturization will New applications stimulate the need for new solve this problem and launch booster cost will materials, and new materials, in turn, create then be minimal (25-35% of the current cost new applications. Future applications for level of a single satellite). GFRP are mirrors of all sizes, surface figure This same idea (cheaper, faster, better) can and accuracies in the IR and visible range, also apply to aircraft equipment and instrularge antenna reflectors (9.1-27.4 m diameter), ments, namely, airborne avionics, tracking electronic chassis and cardguides, plated RF and targeting instruments, and phased components (MUX Cavities, Diplexers), arrays. Today, a military or commercial airphased arrays, to name a few. The list is oncraft maker may not be able to sell a new going but it is hard to tell whether new GFRP aircraft but he could redesign the avionics materials bring about these applications or the (lightweight) to improve the performance of applications bring about the material. For that aircraft (including helicopters). instance, Amoco’s KllOOX fiber is being experimented with for electronic chassis, but the original demand was for high thermal conduc- 47.5 CONCLUSIONS tive material for heat sinks (thermal straps). Ln The technology of composite materials, espeanother example, thin prepregs were develcially GFRP, is constantly evolving to allow oped to make lighter skins for honeycomb increased aerospace application. This panel structures, but in the future mirrors for increased use of GFRP is for both current type visible range optics will also mostly use these applications as well as genuinely new applicathinner prepregs. tions where GFRP was not previously used. The future for composites, in general, looks
References 1021 promising because of the need for aerospace Krumweide, G.C., Brand, R.A. 1991. Attacking Dimensional Instability Problems in equipment and instrument structures to be Graphite/Epoxy Structures. Composite Design, lighter which increases performance and saves Manufacture and Application. ICCM/8. costs. Manufacturing methods are being Honolulu. July 15-19. developed that allow GFRP hardware, in par- Krumweide, G.C. 1977. Development of a graphite/ ticular, to be processed less expensively with Epoxy Reflector: A design-to-cost project. reduced cycle times. It is not hard to imagine SAMPE Quarterly. 8(3). GFRP material replacing most metallic mater- Krumweide, G.C., Chamberlin, D.N., Rule, J.E. 1988.Adaptation and innovation in high-moduial applications for aerospace equipment and lus graphite/epoxy composite design: Notes on instrument structure once 'old paradigms' are Recent Developments. SPIE 0 - E LASIE 1988. given up and awareness of new GFRP materLos Angeles. ial technology is the norm. Krumweide, Derby, E.A., Chamberlin, D.N. 1989. The performance of effective moisture barriers for graphite/epoxy instrument structures. REFERENCES SAMPE. Atlantic City, NJ. Brand, R.A., Derby, EA., Chamberlin, D.N. 1992. Krumweide, G.C., Hoste, J.H. and Staats, J.R. 1979. Structural Development of the Thematic Evaluation of high-modulus pitch/cyanate Mapper Optical Metering Structure. The material systems for dimensionally stable strucEnigma of the Eighties: Environment, tures. Design of Optical Instruments, Orlando, FL: Economics, Energy. S A M P E J. 24(2), 1343-1355. SPIE. Campbell, M.P., Hoste, J.H., Kedward, K.T., Levy, D.J. and Arnold, C.R. 1984. Metal Moisture Barriers for Composites. 29th Nat. SAMPE Krumweide, G.C. 1981. Designing Composite Symp. April 3-5. Structures for Thermally Stable Applications. Fifth DOD/NASA. Conf. Fibrous Composites in Rule, J.E. 1989. Thermal Stability and Surface Accuracy Considerations for Space-based Structural Design. New Orleans, LA. January Single-and-Dual Shell Antenna Reflectors. E S A 1981. ESTEC. Noordwijk, Netherlands. Dodson, DJ., Rule, J.E. 1989. Thermal Stability Considerations for Space Flight Optic Benches. Stumm, J.E., Pynchon, G.E. Krumweide, G.C. 1981. Graphite/Epoxy Materials Characteristics and TomorrowS Materials Today. Vol. 34. Design Techniques for Airborne Instrument Dunbar, D.R., Robertson, A.R., Kenison, R. 1978. Applications. 309. Airborne Reconnaissance. V, Graphite/Epoxy Booms for the Space Shuttle SPIE. Remote Manipulation. ICCM 11, Toronto, Canada. General Dynamics Corporation (GDC Manual). Stumm, J.E., Pynchon, G.E., Pepi, J.W. and Bovenzi, F.G. 1979. Low Temperature/High Stability 1985. Design for Cost and Quality Manual. Applications of Composites. The Teal Ruby Herrick, J.W. Multi-Directional Advance Experiment. Conf. Advanced Composites. El Composites for Improved Damage Tolerance. Segundo, CA. Composites in Manufacturing 3. Anaheim. Telkamp, A.R., and Derby, E.A. 1990. Design January 10-12,1984. Considerations for Composite Materials used in Hertz, J. Moisture Effects on Spacecraft Structures. the Mars Observer Camera. Advances in Optical 1977. The Enigma of the Eighties Environment: Structure Systems. 1303. Orlando, FL: SPIE. Environment, Economics, Energy. SAMPE, 24(2). Kilpatrick, M.C., Girard, J.D., Dodson, K.J. 1990. Walrath, D.E. and Adams, D.F. 1979. Moisture Absorption Analysis of the Thematic Mapper Design of a Precise and Stable Composite Graphite/Epoxy Composite Structure. Modern Telescope Structure for the Ultraviolet Developments in Composite Materials and Coronagraph Spectrometer (UVCS).Advances in Structure. ASME Winter Meeting. Optical Structure Systems. Vol. 1303. Orlando, FL: SPIE.
AIRCRAFT APPLICATIONS
48
Richard N . Hadcock
48.1 INTRODUCTION
Advanced composites, composed of highstrength, high modulus, low density continuous fibers embedded in polymer matrices, first became available some 30 years ago. Since then, composite aircraft structures have transitioned from laboratory curiosities into low-risk, light-weight alternatives to metal structures. Thousands of safety-of-flight composite components are flying in regular service on military and civil aircraft. Major advantages of high-performance composite structures include weight savings, material tailorability, improved fatigue and corrosion resistance. Disadvantages are primarily cost related. Almost all the composite structures currently in production and service have thermoset matrices. A few aircraft parts are currently being made from thermoplastic matrix composites. Metal matrix composites are still in the development stage. 'Conventional' aircraft structural materials now include polymer matrix composites in addition to aluminum, titanium and steel. The chronology of utilization of different reinforced plastics is shown in Fig. 48.1. 48.2 DESIGN AND CERTIFICATION REQUIREMENTS
The general structural design and certification requirements apply to both composite and
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
metal aircraft structures. However, there are some basic differences between the generic civil and military requirements and further variations associated with the intrinsic differences between the structural behavior of metals and composites. The US Federal Aviation Regulations (FAR) for civil aircraft and the US Air Force (USAF) and US Navy (USN) military aircraft requirements have differences. The European Joint Airworthiness Requirements (JAR) are similar to the FAR and have a similar numbering system, but, again, there are some differences. FAR sections related to design and certification of aircraft, aero engines and propellers contain more than 600 pages'. The equivalent USAF documents contain almost 400 pages and distribution is restricted. Although the requirements are applicable to both metal and composite structures, certification of composite structures is very much more extensive and requires many more tests. The contractor must fabricate and test thousands of specimens and hundreds of subcomponents to qualify a single new carbon/epoxy system and associated structural details. As an example, the material qualification program for the all-composite Beech Starship included collection and analysis of thousands of data points from element, panel and subcomponent tests. Finally, fatigue and residual strength tests were made of a complete airframe2f3. A similar test program was used by - CASA to certify the Airbus A320 carbon/epoxy stabilizefl.
Advanced composite materials
Colton/ x- -.- x P r o p e l l e r s
Composite M a t e r i a l
x--RhD.-x
1023
Development
8.lUllt.
FbuPhenolk x - - - x
Winp Spar, Fuselage
Produdlon
Fig. 48.1 Aircraft structural reinforced plastics. (01994 RNH Associates.)
Certification of military aircraft composite structures also requires a major test program. The ’Building Block‘ approach, which starts with static, fatigue and damage tolerance tests of elements, followed by tests of subcomponents and full-scale components, has been very successful in identifying and solving design and manufacturing problems before the design is finalized and frozen. Live fire tests are also required for US military aircraft certification. The ’Building Block’ approach was for certification of the Grumman/US Air Force B-1A composite horizontal stabilizer and for the McDonnell Douglas/US Navy F-18 and AV-8B composite wings5r6.
48.3 ADVANCER COMPOSITE MATERIALS
Boron fiber reinforced plastics (BFRP) and carbon fiber reinforced plastics (CFRP) advanced composite materials were first used for aircraft structures in 1966. Epoxy matrices were used for these materials, which were produced as prepregs in the form of unidirectional prepreg tape. Use of BFRP has been limited by the very high material price (more than $900/kg in 1993). Since the early 1970s, CFRP prepregs have become available in the form of unidirectional tape, woven broadgoods and other forms with many different types of carbon fibers and with many different epoxy, polyimide and bismaleimide (BMI) thermoset matrices. CFRPs have also been available with PEEK and other thermoplastic (TP) matrices.
1024 Aircraft applications Because the price of aviation fuel has dropped dramatically since 1978 relative to material and labor costs, aircraft prices have now become far more important a design requirement than weight savings. As such, it should now be assumed that the customer will pay very little premium for weight savings and the prices of composite components should be comparable to their metal counterparts. Early in the design phase, weight savings can be transformed into cost savings by reconfiguring the airplane, but this is not possible once the overall design is frozen. Weight savings are most valuable early in the design phase, but the value of weight savings diminishes as the design becomes finalized. The design of the structure is primarily dictated by production costs and by maintainability and repair considerations. Composite material prices are very high and costs of design, certification testing, tooling, inspection, material storage, waste material disposal and repair are all higher than for aluminum structures. However, these higher costs can be offset by designing large integral composite components which reduce part count, the number of joints and assembly costs. Most aerospace companies are using the 'Concurrent Engineering' approach, where engineering, manufacturing, quality control, logistics support and cost estimating personnel are formed into collocated teams. The more important interfaces are listed in Fig. 48.2. The costs and prices of aircraft components vary considerably with the type of aircraft and the type, construction, size and materials used for the component. Current (19924) civil aircraft prices vary from $70/kg ($30/lb) for a small private aircraft to between $800/kg and $lOOO/kg for 48.4 THE DESIGN PROCESS airliners and business aircraft. Military aircraft Composites Structure design involves many costs are very much higher and are very different disciplines and is far more complex dependent on the total numbers of aircraft prothan metal structures design because of the duced. These range from a cumulative average anisotropic strength and behavior of the com- cost of $2100/kg for 3000 General Dynamic posite material7. F-16 fighters to more than $12 OOO/kg for 20
Almost all CFRP aircraft structural components have been made with thermoset matrices (1993 prices ranged from $60/kg for epoxy matrix unidirectional tape to $250/kg for BMI matrix materials). Thermoplastic matrix applications have been limited by the high price (about $250/kg) and low compression strength of these materials. Glass fiber reinforced plastics (GFRP) consisting of E-glass, and the higher strength S-glass fibers in epoxy, polyimide and phenolic matrices are being used for helicopter rotor blades, for many secondary structures such as radomes and fairings. Glass/phenolic thermoset and glass/polyethersulfone and other thermoplastic matrix composites are used for fire-resistant cabin and freight hold panels and liners because of their low heat and smoke release properties. Prices of glass /epoxy prepreg range from $5/kg to $10/kg. DuPont introduced KevlaP 49 aramid fiber in 1971. Aramid fiber reinforced plastics (AFRP) have low density, high tension strength and excellent impact damage resistance, but low compression strength. AFRP materials are being used for radomes, leading edges, fairings, floors and other secondary structure applications. AFRP has also been used with some local CFRP reinforcement for the complete airframe of the Avtek 400A business aircraft. Most of the composite primary and safetyof-flight structure components (wings, fuselages, empennage and control surfaces) currently in service are made from carbon/epoxy prepreg unidirectional tape or woven broadgoods. Some light aircraft are made almost entirely from GFRP woven broadgoods.
The design process Requirements
1025
I
External geometry Interface geometry External loads Structural requirements Interface requirements Operational environment Weight & cost targets
Design engineering Internal loads (from stress) Materials selection (with M&P and stress) Cost/weight tradeoffs Detail design Joints and attachments Detail drawings (Paper/computer)
I Engineering
Production
Loads & dynamics Loads Stiffness requirements Flutter analysis, etc.
Materials & processes Material specifications Process specifications M&P standards
Stress Finite element models Design allowables (with M&P) Structural optimization (with weights) Stress analysis Test requirements Stress reports
Manufacturing Manufacturing engineering Manufacturing methods Shop instructions Subassembly Assembly
Weights Weight targets Weight control Structural test Element tests Subcomponent tests Static and fatigue tests Environmental tests Flight test support
Other Quality control Material control NDI equipment NDI/DI requirements Accept/reject criteria (with stress and M&P) Dimensional inspection MRR action/damage repair (with stress, design and M&P)
Facilities/equipment Autoclaves/presses/ovens Refrigerators ATMs/TPMs NC ply cutters RTM/RIM equipment Machining/ drilling/ trimming
Cost estimatingkontrol Manufacturing Materials (including costs of waste and waste disposal) Engineering Quality control MRR actions IE, facilities and equipment Rates and overheads
Tooling Tool design Moldform tooling Bonding tooling Assembly tooling
Product support Maintenance & repair manu LS Inspection requirements Repair mater& and processes Special repairs
Fig. 48.2 Design/manufacturing interfaces. (0 1994 RNH Associates.)
i”’
1026 Aircraft applications
1000 100
MILITARY TRANSPORTS
v)
Z
AIRLINERS
1
BUSINESSNC
7 tff w-
0
PRIVATE AIC
1
U
a
0.1
0.01 1000
a
1000000 10000 100000 OPERATING EMPTY WEIGHT, kg
CIVIL v)
Z
0 i
1
7
te W-
0 a a
b
OPERATING EMPTY WEIGHT, kg
Fig. 48.3 Aircraft costs and prices. (a) Fixed wing aircraft; (b) helicopters. (01994 RNH Associates.)
Structural applications 1027 48.5 STRUCTURAL APPLICATIONS
During the past 30 years, structural applications of composite materials have grown from a few radomes and fairings to the entire airframe. Many hundreds of CFRP wings, stabilizers and flight control surfaces are currently flying in regular service. The composites weight fraction has also
grown during the past 30 years. Published weight fractions are based on either airframe weight, structure weight (airframeplus landing gear), or operating empty weight; and in some cases the base is not defined at all. Because of these differences, the structural composite material weight fractions shown in Fig. 48.4 have been normalized in terms of per cent structure weight (airframe plus landing gear).
40-
xi----------
AV-88
Rafale DGripen 1
30
\
\
I
25
20 15 10
0 YEAR (FIRST FLIGHT)
Fig. 48.4 Composite material weights. (a) Tactical aircraft; (b) civil transports. (01994 RNH Associates.)
1028 Aircraft applications The composites weight fraction for tactical aircraft, Fig. 48.4(a), appears to have leveled between 22% and 26%. The Lockheed F-22A airframe is composed of 26% composites, 30% titanium and 14% aluminum; the McDonnell Douglas F/A-l8E structural weight is 22% composites, 15%titanium, 29% aluminum and 14%steeP. Composite weight fractions for other production and technology demonstrator military aircraft, such as the Grumman A-6E (which has a CFRP wing designed and manufactured by Boeing), Northrop B-2A bomber and YF23A fighter, General Dynamics F-16XL, Grumman X-29A, Rockwell/MBB X-31A, IAI Lavi, the British Aerospace EAP, the Eurofighter 2000 and the Mitsubishi FS-X, are not included because of space limitations or availability of weight data. The composites weight fraction for civil transports, Fig. 48.4(b), has two branches: the
upper branch includes most of the Western European airliners and the Boeing 777; the lower branch includes the McDonnell MD-11, Ilushin 11-96 and the Boeing 737X. The latter aircraft is reported to have less composites than the Boeing 737-300 because many of the operators will be the smaller airlines which do not have composite maintenance and repair facilities12. Weight savings provided by composites vary considerably with the type of aircraft and component. Weight savings, in terms of composite weight fraction, are shown in Fig. 48.5. These tend to decrease as the overall composite weight fraction increases. These data were obtained from different sources, primarily References 7, 9 and 12-18. It should be noted that the weight savings for the AV-8B are based on a weight estimate for comparable aluminum wing, and not the AV-8A aluminum wing, which is smaller15.
A TACTICAL AIRCRAFT I
TRANSPORT AIRCRAFT
I-
10
20
30
COMPOSITES, % STRUCTURE WEIGHT
Fig. 48.5 Composite weight savings. (01994 RNH Associates.)
Structural components 1029 48.6 STRUCTURAL COMPONENTS
Many composite components are currently being produced in the US and overseas for military and civil aircraft. Most of the major commercial aircraft primes subcontract out the production of more composite components than they build in-house. Composite structural and semi-structural components which are currently being produced are listed in Table 48.1. Some specific aircraft components are described later in this section. Different forms of construction have been used for many of these components. Some
combat and technology demonstrator aircraft wings have composite skins which are bolted to metal substructure (e.g. the McDonnell Douglas F/A-18, Grumman X-29, Rockwell X-31, General Dynamics F-l6XL, IAI Lavi), other wings have composite skins which are bolted to composite substructure (e.g. McDonnell Douglas AV-8B, Bell/Boeing V-22, SAAB Gripen, Grumman/Boeing A-6E, Dassault Rafale, British Aerospace EAP, Eurofighter 2000)6,9,12,14,15,17. At least three different forms of construction have been used for stabilizers: composite skins bonded to metal/honeycomb substructure (Grumman F-14, McDonnell Douglas F-15
Table 48.1 Aircraft composite structural components Wings:
Fuselages:
Propulsion systems:
Box beam skins Box beam substructure Winglets Leading edge flaps/slats Ailerons/Flaperons Flaps Flap vanes Spoilers Fixed leading edges Fixed trailing edge panels Flap track fairings Actuator fairinns
Radomes Forward fuselage Canopy frames (helicopters) Mid fuselage Rear fuselage Speedbrakes Tailcone Floor beams Floors Rotordomes Crew compartment armor Rotary launchers Cabin doors (helicopters) Cabin and freight hold linings and partitions Overhead baggage compartments Air ducts Flight refuelling probes
Engine fan blades Engine casings Nozzle flaps Thrust reversers Engine nacelles and cowlings Fan cowls Engine doors Fences Transcowls Firewalls Turbine blade containment rings Drive shafts Pylon fairings Fuel tanks Drop tanks Propeller blades
Empennage:
Horizontal stabilizerslcanards: Skins Substructure Elevators Leading edges Fixed trailing edge panels Tips Vertical stabilizers:
Skins Substructure Rudders Leading edges Fixed trailing edge panels Ventral fins Tips Antenna housings
He1icopter/Rotorcraft: Main rotor blades
Tail rotor blades Rotor drive shafts Doors and fairings:
Landing gear doors Landing gear fairings Landing gear pods Wing-fuselage fairings Stabilizer fairings EauiDment access doors
Miscellaneous:
Crew seats Toilet waste tanks Filament-wound air bottles Cargo containers Cabin door balance springs Tooling - moldforms - assembly/bonding fixtures - checking fixtures Honeycomb core (GFRP and CFRP) Damage repair (metal and composites)
1030 Aircraft applications and F-18, Mikoyan Mig-29), composite skins stabilizer of the Grumman F-l4A, Fig. 48.6. bolted to metal substructure (General The Navy requirement that the stabilizer Dynamics F-16), and composite skins bolted to should be fully qualified by full-scale static composite substructure (McDonnell Douglas and fatigue tests one year before the first flight AV-8B, Bell/Boeing V-22)9,15*17,20J1. was satisfied by successful tests in 1969. More Elevators, rudders, ailerons and other con- than 1500 stabilizers had been built when F-14 trol surfaces have generally been made by production ended in 1991l'. bonding composite skins to substructure com-
spars. All composite sheet-stiffened construction is used for the forward fuselage of the McDonnell Douglas AV-8B and the complete fuselage of the Bell/Boeing V-229J4,16. There are also many variations in the construction of composite components used for civil aircraft and helicopters. 48.6.1 MILITARY AIRCRAFT COMPONENTS Development of advanced composites technology was initiated in 1964 by General Bernard Schriever of the US Air Force who tasked the Air Force Materials Laboratory (AFML) at Wright Patterson Air Force Base, Ohio, to contract some US aircraft companies to design, fabricate and test various structural components made from boron/epoxy. McDonnell made 45 rudders which were installed on F-4 fighters for service evaluation; General Dynamics was awarded a major contract to develop a boron/epoxy horizontal stabilizer for the F-111 in 1966; and Grumman won a contract to design and fabricate a wing box beam for a conceptual advanced fighter in 1968. The US Navy awarded a contract to Douglas to develop a carbon/epoxy horizontal stabilizer for the A 4 in 1971. Douglas had previously made and flown a carbon/epoxy wing flap on an A-4. By May 1971, more than 80 different advanced composite aircraft component programs had been initiatedzJ3. The first production application of advanced composites for a safety-of-flight component was the boron/epoxy horizontal
Fig. 48.6 (a) Grumman F-14; (b) F-14 borodepoxy horizontal stabiliser.The stripes on the lower cover are aluminium foil for lightning protection. (Courtesyof Grumman Corporation.)
Boron/epoxy covers were utilized for the McDonnell F-15 horizontal and vertical stabilizers and rudders. The large speed brake, which is located on the fuselage behind the canopy, is carbon/epoxy-aluminum honeycomb sandwich. The F-15 was still being produced in 1994,20 years after the first flight. The principal production use of boron/epoxy was for the F-14 and F-15 stabilizer covers. Boron/epoxy was selected
Structural components 1031 because a comprehensive design database was available for the material from Air Force development programs and was less expensive than carbon/epoxy. During the intervening years, this situation has changed. Extensive design data have been generated for many different carbon/epoxy systems and prices are now significantly lower than for boron/epoxy, Carbonlepoxy was selected for the horizontal and vertical stabilizers covers of the General Dynamics F-16. These covers are bolted to aluminum substructure. More than 3000 F-16s have been produced between since 1976. McDonnell Douglas, with Northrop as the major subcontractor, was awarded the Navy Air Combat Fighter contract in 1976 with the F/A-l8A, an enlarged and extensively modified version of the Northrop YF-17. To save weight and offset the increased weight of the landing gear and other components required for carrier operation, carbon/epoxy materials were used for the wing, tail, control surfaces
and access panels. The aircraft is currently being extensively redesigned and enlarged. Designated the F/ A-1 8E / F, carbon / epoxy accounts for 18%of the structure weight compared to 10% for the F/A-18A. Because of its higher strength and stiffness, the new IM7/977-3 carbon/epoxy system is being used for the wing and empennage skins instead of AS4/3501-6’J2. F/A-l8E/F material usage is shown in Fig. 48.721. In 197677, McDonnell Douglas extensively redesigned the British Aerospace AV-8A Harrier vertical take off and landing (VTOL) fighter/attack aircraft for the US Marine Corps to improve range/payload capability. Carbon/epoxy was utilized for the complete wing, horizontal stabilizer and forward fuselage. Designated the AV-8B Harrier 11, the composite wing weight is about the same as the aluminum wing of the AV-8A but has 19%more area, a 50% increase in internal fuel capacity and a supercritical airfoil. Range-payload capability was increased by 100%. Because of high temperatures from the exhaust in the VTOL
Percent of Structural Weight UAlurninurn
FIA-18CID FlA-18UF 49 ________ 31 15 14
HSteel ___________ ________ NUilTitaniurn 13 ________ 22 Carbon EDOXV.--10
________
18
Fig. 48.7 McDonnell Douglas F/A-l8E/F material usage. (Courtesy of McDonnell Douglas.)
1032 Aircraft applications mode, carbon/bismaleimide is used for the skins of the inboard flaps and the strakes, which are mounted under the fuselage9J2. Usage of composites, which account for 26% of the airframe weight, is shown in Fig. 48.8.2l Composites are also used on other US tactical aircraft. The Navy funded Boeing to design a CFRP wing for the Grumman A-6E in 1987 because fatigue of the aluminum wings limited the life of the aircraft. Boeing built 179 composite wings in Seattle, which were installed on the last batches of 21 A-6 aircraft. An additional 158 aircraft were retrofitted with CFRP wings. A follow-on program for 120 wings was canceled by the Navy in September 1993. The first production application of carbon/thermoplastic composites was the stabilizers of the Lockheed F-117A interdictor. The complete fleet of F-117A aircraft are being filled with these stabilizers to extend the flight envelope9J2. The wing, fuselage, tail and the rotors of the Bell/Boeing V-22 Osprey multi-mission VTOL tilt-rotor aircraft are almost entirely made from CFRP and GFRP. Composites, which
account for 44% and 7% respectively of the structure and rotor weight of the aircraft, are estimated to have provided a 25% weight saving relative to a comparable metal airframe1*J4. Because weight is so critical to performance of this aircraft, much of the structure is currently being redesigned to further reduce weight and costs'*. The outer skin and much of the substructure of the large Northrop B-2A bomber is made from carbon/epoxy. The B-2A was designed and produced by a team composed of Northrop, Vought and Boeing. The B-2A, which has a wing span of 52 m (172 ft), a wing area of 464 m2 (5000 ft') and a gross weight of 180 metric tomes (4000001b), first flew on July 17,1989 and is by far the largest 'composite' aircraft to fly to date9J2. McDonnell Douglas is using 6800 kg (15 000 lb) of composite materials for control surfaces, stabilizer leading edges, the tail cone, landing gear doors and pod, engine nacelles, and the wing-fuselage fairing of the US Air Force C-17A transport. These components, shown shaded in Fig. 48.9, account for 8% of the structure weight
Fig. 48.8 McDonnell Douglas/BAe AV-8B material usage. (Courtesy of McDonnell Douglas.)
Structural components 1033 and are made by McDonnell Douglas and many different American and European subcontractors including Grumman (ailerons, elevators, rudders), Heath Techna (fairings), Beechcraft (landing gear doors), Aerostructures Hamble (flap hinge fairings and trailing edge panels) and Northwest Composites (main landing gear pods). Since 1970, more than 7500 fixed wing US military aircraft have been produced which utilized composites for safety-of-flight components, such as horizontal stabilizers, wings and control surfacess. Composite materials have also been used for many structural components of European and other military aircraft. In France, Avion Marcel Dassault-Breguet Aviation and Aerospatiale have been involved in composite structures development programs with the French government support since 1972. A boron/epoxy rudder was flown on a Dassault Mirage I11 in 1975; horizontal stabilizers were flown on the Mirage F I in 1976 and composites first entered production in 1978 with CFRP ailerons for the Mirage F I, which provided
26% weight saving26. The Dassault Mirage 2000, which first flew in 1978, has a CFRP fin designed and built by Aerospatiale, a CFRP/BFRP hybrid rudder, and CFRP elevons, nose landing gear doors and some equipment doors. Composites account for about 12% of the structure weight and provide a 25% weight savings of that total. About 350 Mirage 2000 had been delivered by the end of 1993133. Dassault Rafale C and M tactical combat aircraft are now in production with CFRP wings, forward fuselage, vertical stabilizer, fin, and control surfaces with an AFRP radome and jet pipe fairings, accounting for about 25% of the structure. Other advanced materials include superplastic formed-diffusion bonded titanium leading edge flaps and canards and aluminum-lithium fuselage panels9J3. British Aerospace (BAe) designed and built a CFRP wing for the SEPECAT Jaguar and teamed with MBB, in Germany, to design, build and fly a CFRP taileron for the Panavia Tornado in the early 1980s. The British government authorized BAe to go-ahead by with
NONSTRUCTURALPARTS, LINERS, TROOP SEATS
7
. . GFRPlNomex Core KevladFoam Core Kevlar/Nomex
Fig. 48.9 McDonnellDouglas C-17A composite applications.(Unpublished:all rights reserved under copyright law by McDonnell Douglas. Reprinted with permission.)
1034 Aircraft applications development of the EAP (Experimental Aircraft Programme) in 1982. The EAP had a carbon/epoxy wing with carbon/epoxy substructure co-bonded to the lower cover. BAe made the right wing and Aeritalia, who contributed about 15% of the costs of the program, made the left wing9J3.The EAP flew in 1986 and the CFRP wings provided much of the technology and background of experience used currently by BAe Aeritalia (now Alenia), CASA and MBB-Dornier (now DASA) for Eurofighter 2000 (EFA) carbon/epoxy wings, fin,rudder and control surfaces. The first EFA flew in 19939. The EAP wing also provided BAe with the experience to assist SAAB in the design of the CFRP wing for the Swedish JAS 39 Gripen multi-mission fighter. The first three and a half wing sets were manufactured by BAe, subsequent wings and all other CFRP components are manufactured by SAAB in Sweden. CFRP components include the wings, canards, fin, outboard elevons, landing gear doors, fuselage and fin fairings and some fuselage access doors, shown in Fig. 48.10. The Gripen first flew in 1988 and entered service with the Swedish Air Force in 19939,'3J7. Other military aircraft which have AIRFRAME advanced composite components and are curThe airframe S ~ N C ~ I Cis~deslgned for hlgh durablilty and lo modern damage tolerancecdteda. To keep the welght low, abaut 25 %of the rently being produced include the Russian s t i c h m Is made lmm carbon flbre CompoGltes. Mikoyan MiG-29 fighter and the Ukrainian Antonov An-124 heavy transport. The rear portions of the horizontal stabilizers, fins, mdders, ailerons and flaps of the MiG-29 have carbon/epoxy covers, which account for about 7% of the airframe weight. Mixed composites are used for many secondary structure applications on the An-124 stabilizer including trailing edge panels, landing gear pods, wingfuselage and flap track fairings, floor beam caps, and nacelles. About 5500 kg (12 125 lb) of (c) Carbon Fibre CompositesContent CFRP, AFRP and GFRP are used on the An124, saving 1800 kg (3968 lb)9,13. Many technology development and flight demonstration programs which have incorpo- Fig. 48.10 SAAB JAS39 Gripen (a) Composite wing rated CFRP wings and other components. assembly; (b), (c) composite applications. (Courtesy These include the Rockwell International of SAAB Military Aircraft.)
Structural components 1035 HiMAT, Grumman X-29A, Israeli Aircraft structures. Lockheed designed and made 18 Industries Lavi (the CFRP wing and vertical AFRP fairing panels and 8 AFRP ailerons for stabilizer were designed and built by L-1011 Tristars. Lockheed changed many Grumman), the General Dynamics F-16XL and glass/epoxy panels and fairings to the Rockwell International/MBB X-31A9. Kevlar /epoxy for the extended-rangeTristars3I. Aeroelastic tailoring was used on the The experiences gained from the ACEE proHiMAT and X-29A to improve maneuver capa- grams provided the confidence needed by bility and, in the case of the X-29A, prevent Boeing to select CFRP for the Boeing 757, 767 divergence of the forward-swept wing2s,29. and 737-300 control surfaces in the late 1970s. Dornier designed, built and ground tested a Boeing uses about 1500 kg (3300 lb) of CFRP wing for the Alpha Jet in 198630. advanced composites in the 757 and 767 and 680 kg (1500 lb) in the 737-300. CFRP components include elevators, rudders and many 48.6.2 CIVIL AIRCRAFT APPLICATIONS secondary structure components made from Prior to 1972, the major US aerospace compa- CFRP, GFRP and CFRP/GFRP hybrids. By the nies advanced composite development end of 1988, composite components on Boeing programs were primarily directed toward mil- airplanes had accumulated more than 9-military applications of these materials. Boeing lion flight hours3*.Most of the 757 and 767 flew a boron/epoxy fore-flap on a 707 in 1970, fairings and fixed panels were originally but commercial aircraft advanced composites aramid / epoxy or aramid / carbon/epoxy development essentially began when the hybrid-honeycomb sandwich construction. ACEE (Aircraft Energy Efficiency) Program Because of surface and matrix cracking, much was initiated by NASA in 1972. The program of the aramid/epoxy was replaced with determined properties of different carbon/ glass/epoxy, to improve surface p r ~ t e c t i o n ~ ~ . epoxy material systems after long-term worldFollowing definition in 1990 of the 777, the wide environmental exposure and spectrum world’s largest twin-engined jet, Boeing fatigue loading equivalent to 20 years or 36 000 formed design/build teams to develop the airflights of airline service. The program also frame and systems. About 15 000 kg included flight service of numerous composite (33 000 lb) of composite materials are used on components to obtain confidence in the long- the 777. The empennage, produced by Boeing term durability of advanced composite at the new Composites Manufacturing Center at Fredrickson, Washington, is made from a structures and materials3*. The Boeing carbon/epoxy flight service new toughened-matrix carbon/epoxy comprogram, which began in 1973, included 108 posite prepreg, manufactured by Toray in a Boeing 737 spoilers, ten 727 elevators and four plant next to the Fredrickson facility. The same 737 horizontal stabilizers. The McDonnell material is used for the floor beams, which are Douglas carbon/epoxy program included 20 produced by Rockwell International. Other DC-10 carbon/epoxy rudders and a single composite components are supplied by comvertical stabilizer. By 1991, one of the DC-10 panies in the USA, Canada, Europe, Australia, rudders and the vertical stabilizer, which was Singapore, Brazil and Korea. The first 777 flew installed on a Finnair DC-10, had respectively in June 1994 and deliveries began in July 1995. accumulated 58 300 and 17 580 h o ~ r s l * , ~As ~ .of September 1996, Boeing had delivered 35 Lockheed made a carbon/epoxy vertical stabi- aircraft and had 281 orders from fourteen airlizer for the L-1011, which was tested but not lines8J2. Composite structure on the 777 is shown in Fig. 48.11. flown. The stabilizers have laminated skins with The Lockheed flight service program was primarily directed to Kevlar/epoxy (AFRP) cocured stringers, solid laminate spars, simple
1036 Aircraft applications m composite materlais Toughenedmaterial for Improved damage reslstaneaand damagetolerance, and parts are deslgnedfor slmple, bolted or bondedIwairS Corroslon andfatlgue reslsmt * Welght savings
-
ITwghened graphite I Glsphlte 1 HYbm
Outboard alleron fln torque box Stablllzertorque box
6 Flb.wb Wlng fixed leading Trailing edge panel
leading and trailing
Nose radome
Fig. 48.11 Boeing 777 composite structure. (Courtesyof the Boeing Company.)
honeycomb sandwich ribs and non-structural forward torque boxes. They are designed for simple repair. The horizontal stabilizer has a span of 21.3m (70 ft) and the projected tail area is 101 m2 (1090 ft2),about twice the span and four-times the area of the AV-8B ~ i n g ~ Following successful airline service of the DC-10 CFRP rudders and fin, deployed under the ACEE program, and production experience from composite components on the MD-80, McDonnell Douglas extended use of composites on the MD-11 wide-body airlineP. Almost 5000 kg (11000 lb) of composite materials are used on the MD-11 of which 4400 kg (9700 lb) is structural (primarily carbon/epoxy). Weight savings of 20-30% have been achieved after trading some weight savings for improved reliability, maintainability, durability and producibility. Most of the composite structure is solid-skin or Rohracell foam sandwich construction. Honeycomb sandwich construction has been avoided because of susceptability to impact damage and moisture penetration into the core1*.MD-11 composite components are shown in Fig. 48.12. As a result of the very successful French Dassault-Aerospatiale composites develop-
ment programs, which included components for military aircraft and a complete CFRP wing for the Falcon 10, Airbus Industrie began using carbon/epoxy for the A-300-600 spoilers in 1983 and for the complete vertical stabilizer, , rudder, ~ ~ . elevators and spoilers of the A310-300 in 1985. The vertical stabilizer is used as a fuel tank on extended range aircraft26. The entire tail, control surfaces, and cabin floor of the A320, A321, A330 and A340 are carbon/epoxy. The A300 medium-range narrowbody airliner entered airline service in 1988 the larger A330 and A340 wide-body airliners entered service in 1993 and 1991 respectively. Composite applications on the A320 are shown in Fig. 48.1326. The Airbus CFRP horizontal stabilizers are designed and manufactured by CASA in Spain and the vertical stabilizers are manufactured by Deutsche Airbus in germ an^^,^^. Avions de Transportation Regionale (ATR), the Aerospatiale/ Alenia consortium, uses carbon/epoxy for all the control surfaces and Kevlar/epoxy for many components of the ATR 42 regional airliner which was certificated in 1985.
Structural components 1037 Winglet Trailing Edge
No. 2 Engine inlet Access Panels
Outboard Ailerons
Maln Landing Gear Door and Struts OF-*
Fig. 48.12 McDonnell Douglas MD-11 composite structure. (01991 by McDonnell Douglas Corporation. All rights reserved. Reprinted by permission.)
ATR utilizes carbon/epoxy for the complete outer wing boxes of the ATR 72, a stretched version of the ATR 42, the first airliner with a CFRP wing box to be fully certificated in Europe and the USA. It first flew in 1988 and was certificated and entered airline service in 1989. The wing box, which is a fuel tank, is made by Aerospatiale at Nantes. The weight of the wing box was reduced by 130 kg (286 lb) using CFRP instead of a l u m i n ~ m ' ~ , ~ ~ . Composite materials used on the ATR 72 and details of the CFRP wing box are shown in Fig. 48.14. Domier chose CFRP for the complete tail and rear fuselage of the Domier 328 regional airliner, which also has an AFRP pressure bulkhead. The 328 was fully certificated and entered airline service in 1993. Material distribution and details of the CFRP rear fuselage/fh box structure are shown in Fig. 48.15. The de Havilland Canada Dash 8, Embraer Brizilia, SAAB 340 and other regional airliners are using AFRP or CFRP for many secondary structure components.
Many thousands of structural components made from composites are currently in airline service. Damage assessment and repair has been a major problem to the airlines. Too many different fiber-matrix systems are currently being used, even by the same OEM; prepreg materials are expensive and have a limited life; damage assessment requires special nondestructive test equipment and experienced technicians; minor repairs need special skilled mechanics; major repairs must be made in an autoclave and take many days to complete; cost of replacement parts or leasing spares are much higher than metal part^^^,^^. IATA sent a questionnaire to the airlines to obtain information on composites maintainability in December 1991. The ATA/IATA/SAE Commercial Aircraft Composite Repair Committee (CACRC) was formed the following year. Committee members include representatives from the FAA, airlines, OEMs, NASA and material s ~ p p l i e r s ~ ~ .
1038 Aircraft applications
m CFRP
Ailerons
I AFRP
CFRPIAFRP GFRP
Spoile shroud
j H o r c o n t al Iize r s stabi . and elevators
Landing gear doors
trailing edge access panels
Fairings
Fig. 48.13 Airbus A320 composite applications. corbcn
1 nomex sandwich
c o k n rnonolimic structure
A
dsm
kwhr 1 n o m u sandwich kevlar 1 nornex sandwich wifh stiffening carbon plies fib-lair I nomex sandwich
A
cobin floor an& : kevlar / nomex sandwich propeller plader : fibreglorr / polyurethane foorn / d ~ r n i n i ~spar m broker : carbon
5,.Ib,II
CARBON SPARS (FRONT AND RFAR )
(b) CARBON PANELS
(TOP AND BOTTOM)
Fig. 48.14 ATR72 (a) composite applications; (b) CFRP wing box. (Courtesyof ATR.)
91
Structural components 1039
Domier 328
CFC Components of the Rear Fuselage and the Vertical Stabllirer Box
653515.11 1,~ , ,u ,
~
(b) Fig. 48.15 Domier 328 (a) composite applications; (b) rear fuselage and fin. (Courtesy of Domier.) 48.6.3 GENERAL AVIATION APPLICATIONS
The Starship and the one-piece CFRP wing, The Windecker Eagle, which flew in 1967, was which has a span of 16.6 m (54 ft) are shown in the first all-composite (GFRP) airplane to Fig. 48.16. The complete rear fuselage, tail, and canard obtain full FAA certification. The Eagle was of the Italian Rinaldo Piaggio P.180 Avanti is followed by the Lear Avia Lear Fan 2100, carbon/epoxy; the Avanti was granted full which had an airframe made from CFRP with in 1990, but only 20 aircraft had certification some AFFW secondary ~tructure'~. The Lear been sold by the end of 1993. Piaggio was Fan program was terminated in 1985 because planning to build 10 aircraft in 1994. of certification problems and costs. Most of the airframe of the Avtek 400A is The Beech Starship 2000, which is almost Kevlar/epoxy with carbon/epoxy reinforceentirely made from carbon/epoxy-faced honment. Avtek was initiating the FAA eycomb sandwich construction, received full certification program in 199338. FAA certification in 198837.By December 1993, Some European all-composite (primarily Beech had produced 50 Starships and producGFRP with carbon/epoxy reinforcement) tion was put on hold because of slow sales.
1040 Aircraft applications
Fig.48.16 (a) Beech starship 2100; (b) starship wing upper cover. (Courtesyof Raytheon Aircraft Co.)
aircraft include the British Slingsby T67 Firefly trainer (USAF T-3A), the German Grob G115 and FFT Eurotrainer. The Russian Sukhoi Su26 and Su-31 aerobatic competition aircraft have carbon/epoxy wings and tails. CFRP and GFRP materials have been used for many experimental and home-built and kit aircraft. These include the Voyager, designed by Burt Rutan and built by Scaled Composites, which made the first unrefuelled non-stop flight around the world in December 1986. 48.6.4 HELICOPTER APPLICATIONS
Glass/epoxy main helicopter rotor blades were initially developed in the late 1960s by
Messerschmitt-Elkow-Blom(MBB) in Germany, Sud-Aviation (later Aerospatiale) in France and Kaman in the United States. The MBB 80.105, which has GFRP blades, was first produced in 1967. With the exception of Sikorsky, almost all of the helicopters currently in production worldwide have composite blades. Aerospatiale (now part of Eurocopter International, which was formed in 1991 with the merger of the helicopter divisions of Aerospatiale and DASA/MBB) began GFRP composites in the mid-1950s for cowls and fairings. About 25% of the structure weight of the AS 365 is made from composites26. Sikorsky used GFRP for the canopy frame of the S-61 helicopter in 1959, and by 1967 composites use had been extended to the rear fuselage skins, doors, the horizontal stabilizer and the engine cowlings of the Sikorsky S-76. Sikorsky and Bell produced composite components for the ACEE program and complete fuselages for US Army Advanced Composite Aircraft Program (ACAP). Composites are used extensively for the rotor blades and much of the airframes of the new Boeing/Sikorsky RAH-66 Comanche. Compared to the UH-60 Black Hawk, composite usage in the airframe has grown from 9% to 51%; has shifted from 50% titanium to 67% composites for the rotor system; and has shifted from 22% magnesium to 28% composites for the drive system. Carbon/epoxy accounts for 44% of the airframe weight; the rotor blades are made from carbon/epoxyglass/epoxy broadgoods; the main gear box housings are made from carbon/bismaleimide using resin transfer molding (RTM);and the tail drive shaft is filament-wound carbon/ epoxf9. Composites are being used extensively for rotor blades and airframes many other new American, Western European and Russian helicopters including the McDonnell Douglas Explorer, Eurocopter BO 108 and the Kamov Ka-62 g.
Structural components 1041 48.6.5 PROPULSION SYSTEM APPLICATIONS
In addition to engine cowlings, cowl doors and nacelles, advanced composites are being utilized for engine casings, thrust reversers, translating cowls, fan blades and propeller blades. A Grumman-developed carbon/bismaleimide aft cowl assembly and thrust reverser for the Fokker 100 is shown in Fig. 48.17(a). This is 25% lighter and far less expensive than the current titanium cowl assembly. A production CFRP transcowl for the General Electric CF6-80C2 engine is shown in Fig. 48.17(b). The transcowl assembly weighs 86 kg (180 Ib). Grumman has been contracted to build 600 units for the CF680C2, which are used on the Boeing 747 and 767, the MD-11 and the A320. 48.6.6 CONCLUSIONS
The weights of composite structures produced annually for airliners and military aircraft, based on five-year averages, are listed in Table 48.2. Twenty years ago, advanced composite structures were only being produced for components of military aircraft. This situation has changed dramatically during the intervening years: the total annual production weight has increased by two orders of magnitude and most of the composite structure (66%)is being produced for European airliners (Airbus, ATR, Dornier).
Fig. 48.17 (a) C F cowl ~ assembly; (b) transcowl. (Courtesy of Grumman Corporation.)
Table 48.2 Composite structure production (five-year averages) Period Total fly weight: metric tonnes/vear Distribution: US military aircraft US airliners Other military aircraft Other airliners
1975-1979
1980-1 984
1985-2989
1990-1994
27
180
905
2620
158 (18%) 334 (37%) 93 (loo/,) 320 (35%)
172 (7%) 582 (22%) 126 (5%) 1740 (66%)
21 (78%) 3 (llY0) 3 (11Yo) -
56 (31%) 97 (54%) 21 (12%) 6 (3%)
1042 Aircraft applications During the past 20 years, advanced cornposite structures have become light-weight, reliable, low-risk alternatives to conventional aluminum and titanium structures. Their use will continue to grow. REFERENCES 1. Federal Aviation Regulations (FAR), Subchapter C - Aircraft. 14 CFR Chapter 1, US Department of Transportation, Federal Aviation Administration, Washington DC. 2. Abbott, Ric, Design and Certification of the AllComposite Airframe. SAE paper 892210, SAE Tech. C o d , September 1989. 3. The All-Composite Airframe Design and Certification.Aerospace Engrzg, April 1990. 4. A. Barrio Cardaba, et al., Design and Fabrication of the Carbon Fiber/Epoxy A-320 Horizontal Tailplane. SAMPE J., Jan./Feb. 1990. 5. Waggoner, Gary and Erbacher, Herman, Damage Tolerance Program for the B-1 Composite Stabilizer. AlAA Paper 77-464, AIAA/ASME 18th Structures, Structural Dynamics and Materials Conf., March 1977. 6. Weinberger, Robert A. et al., US Navy Certification of Composite Wings for the F-18 and Advanced Harrier Aircraft. AlAA Paper 77466, AIAA/ASME 18th Structures, Structural Dynamics and Materials Conference, March 1977. 7. Hadcock, Richard N. Design and Analysis of Advanced Composite Structures. Handbook of Composites, (ed. George Lubin), Chap 20, New York Van Nostrand Reinhold, 1982. 8. Aerospace Facts and Figures 1992-2993(and many previous editions). Aerospace Industries Association of America, Washington, D.C. 9. Jane’s AI1 The World’s Aircraft, 1993-94 (and previous editions), (ed. Mark Lambert). Coulsdon, England: Jane’s Information Group Ltd. 10. Ted G. Nicholas. US Military Aircraft Data Book. Various editions. Data Search Associates. 11. Current Aircraft Prices ’88 (and other editions) Interavia Publishing. 12. Aviation Week and Space Technology: Various issues. 13. Defense News, Flight International, Aerospace, Aeronautics B Astronautics, Aerospace Engineering, Air International: Various Issues. 14. Aerospace Composites B Materials, 1988-1991 and Aerospace Materials, 1992-1993, Burnham, England: The Shephard Press.
15. Schier,J.F. and Juergens,R.J., They Force a Fresh Look at the Design Process. Aeronautics b &ptember 1983. 16. Watson, James C., AV-8B Composite Fuselage Design. Aircraft, March 1982. 17. Lubin, George and Dastin, Samuel J. Aerospace Applications of Composites. Handbook of Cimposites, (ed. George Lubin), Chap 20, New York Van Nostrand Reinhold, 1982. 18. Ashizawa, Moto, Composite Technology Growth Leading to the MD-11 Application and to the Civil Transport Aircraft of Tomorrow. Proc. SAMPE Meeting, Japan, 1991. 19. Brenner, Lothar and Johst, Eberhard, The Airframe of the Dornier 328 - Proven Progress. DGLR Jahrbuch 1989 I. 20. Hadcock, Richard N. and Huber, John. Specific Examples of Aerospace Applications of Composites. Lecture Series No. 124, Advisory Group for Aerospace Research and Development (AGARD), October 1982. 21. Information supplied by the author by McDonnell Douglas Aerospace, St. Louis, Missouri. Reproduced with permission. 22. Rosato, Dominic V. and Lubin, George, Plastics in Aircraft and Aerospace. Handbook of Fiberglass and Advanced Plastics Composites, (ed. George Lubin), Chap 29, New York: Van Nostrand Reinhold, 1969. 23. Hadcock, Richard N., Boron/Epoxy Aircraft Structures. Handbook of Fiberglass and Advanced Plastics Composites, (ed. George Lubin), Chap 24, New York Van Nostrand Reinhold, 1969. 24. Hadcock, Richard N., Status and Analysis of Advanced Composite Aerospace Structures Programs. Composites Technical Note No. CTN466-35. G m a n Aerospace Corp. May 1971. 25. Chaumette, D. Flight Qualification of Composite Structures at Avions Marcel Dassault - BrCguet Aviation. AlAA Paper 820755,1982. 26. Information supplied to the author by Avions de Transport Regional, Blagnac Cedex, France. Used with permission. 27. Information supplied to the author by SAAB Military Aircraft, Scania AB, Linkoping, Sweden. Used with permission. 28. DeAngelis, V. Michael, In-Flight Deflection Measurement of the HiMat Aeroelastically Tailored Wing. J. Aircraft, December 1982. 29. Hadcock, Richard N., X-29 Composite Wing. Evolution of Aircraft/Aerospace Structures and Materials Con5 Air Force Museum/AIAA, April 1985.
References 1043 30. Rose D. et al., Design of a Carbon-FiberCommercial Aircraft Group, Seattle, Reinforced Wing for the Alpha-Jet Major Panel Washington.Used with permission. Tests. 1.Aircraft, June 1986. 35. The Cost of Composites. Air Transport World, 31. ACEE Composite Structures Technology: July 1992. Review of Selected NASA Research on 36. Harris, Charles, Assessment of Practices in Composite Materials and Structures. NASA Supporting Composite Structures in the Conf. Pubn 2321,1984. Current Transport Fleet. 4th NASADoD 32. Harradine, Peter J. and Quinlivan, John T., Advanced Composites Technology Conf., Salt Lake Composites and the Commercial Jet - A Boeing City, Utah, June 1993. Paper No. AIAA-89-2126. 37. Information and photographs supplied to the Viewpoint. AIAA/AHS/ASEE Aircraft Design, Systems author by Beech Aircraft Corporation, Wichita, and Operations Conference, Seattle, 1989. Kansas, 1993. Used with permission. 33. Hyatt, M. Caton, R. and Lovell, D., Advanced 38. Information and photographs supplied to the author by Avtek Corporation, Camarillo Materials Development in Commercial Aircraft. Paper No. AIAA-89-2127. AIAA/AHS/ASEE California, 1993. Used with permission. Aircraft Design, Systems and Operations 39. Garbo, Samuel P. and Rosen, Kenneth M., Composites Usage on the RAH-66 Comanche. Conference, Seattle, 1989. 34. Information supplied to the author by Boeing Vertifite, March/April 1992.
COMPOSITES IN THE SPORTING GOODS INDUSTRY
49
Brian E. Spencer
49.1 INTRODUCTION
In 1990 Frost & Sullivan reported that over $800 million dollars were spent by domestic defense/aerospace contractors for advanced composites. In the same period, $70 million worth of advanced composites were purchased from domestic producers for sporting goods. Advanced composites are usually defined as those composites that use either carbon, aramid, Sglass, ceramic, polyethylene, boron, or other high strength or high stiffness fiber. In 1991, the total worldwide consumption of advanced fibers was estimated to be 27200 metric tons1. Approximately 10% of this fiber
went into sport and recreation applications. Worldwide sport and recreation applications are the third largest user of advanced composites behind defense/aerospace and elastomer reinforcement (tires, hose and belts)'. In the USA as well, sport and recreation applications are the third largest users of advanced composites. One of the major growth markets for advanced composites over the past several years has been the sporting goods industry. Although there was a total decline in USA defense spending of only 14% from 1990 to 1994, the procurement decline was 45%' (Fig. 49.1). The defense procurement decline was
300
",250
Manpower, R&D logistics, other
c ln
=E0 200 c (I)
7 m
Procurement
150
100
50 0
Fiscal year
Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
Fig. 49.1 Defense budget declines. (Courtesy of Westinghouse ESG news and department of defense budget.)
Introduction disastrous for the composites industry and resulted in severe upheavals. Although no segment of the composites market could offset the defense-induced decline, the sporting goods/recreation composite segment along with most other segments is predicted to show positive growth in the near future3(Table 49.1). This growth trend is expected to continue into the 21st century. Advances in materials and processes have reduced consumer prices for the recreational composite while providing improved performance for the athlete. Some of the first applications for composites in sport and recreation were fiberglass boat hulls and fishing poles. Now the list of products using composites includes almost every sporting and recreation activity. Products in golf, tennis, and bicycle racing have brought attention to the superior performance of composites in sports and recreation. Hand lay-up and roll-wrapping have been the processes generally used for most sporting goods applications. There has been an increased interest in filament winding as a preferred process for tubular products such as golf shafts, sail masts, ski poles, softball bats and bicycle frame tubing, because filament winding can lower labor costs and add a new level of design flexibility, product consistency, and quality for these products.
1045
49.2 MANUFACTURING TECHNIQUES
Composite products may be manufactured by several processes, depending on the shape and physical properties required. Common processes are hand lay-up, machine lay-up, roll wrapping, filament winding, pultrusion, compression molding, platen molding, chopped fiber spray-up, resin transfer molding and braiding. Not all of these processes are efficient for manufacturing sport and recreation products. Virtually all fishing poles and most golf shafts are fabricated by the roll wrap process, which consists of placing pre-cut sheets of prepreg unidirectional tape or cloth on a mandrel, then rolling the mandrel between platens to consolidate the composite. Fly fishing rod lay-ups have in the past had all unidirectional 0" plies of one material but golf shafts may have two fibers, such as carbon and boron, and some angle plies. A release cloth or tape is usually wrapped around the composite to apply pressure during the required heat cure process and hold the composite in place. After curing, the composite part is pulled from the mandrel and can be machined as necessary to the final shape. The part is then painted and graphics are added. Braiding and pultrusion have been used successfully for some applications. Tennis
Table 49.1 Annual growth rate of selected materials by region, 1993-2003 _ _ _ _ ~
Productlend use
World average annual growth
US average annual growth %
Rest of world average annual growth YO
6.5 1.0 6.0 6.0 50.0 4.0
8.5 4.0 3.5 4.5 2.0 5.0
7.45 2.2 5.0 5.4 26.7 4.6
7.1
5.4
6.4
%
Advanced polymer composites Aircraft Missiles/ space Recreation Armor Automotive Industrial/other Total
1046 Composites in the sporting goods industry
rackets have been fabricated by braiding. The braiding process for composites is identical to braiding textiles, except that the fiber is usually braided as a prepreg. Prince introduced the Vortex tennis racket, braided with co-mingled nylon and carbon fibers, in the late 1980s. Ski poles, kite tubes, arrow shafts, and similar hollow or solid shafts are made by the pultrusion process. Pultrusion of composites is similar to extruding metals. The resin and fiber material are pulled through a heated die that forms the composite to shape and cures the resin in one continuous process. Since the fibers are essentially all aligned down the axis of the part there is little transverse or hoop strength unless hoop fibers or other off-axis reinforcement such as mats are added during the process. Arrow shafts have shattered on impact because of this lack of strength in the hoop direction. Resin transfer molding (RTM) consists of injecting resin into a mold containing fiber in a preform pattern. Resin transfer molding is usually used in conjunction with a fiber placement technique such as hand lay-up, braiding or filament winding. Some tennis rackets are fabricated using the resin transfer molding process4. The filament winding process is the machine controlled technique of applying a controlled amount of resin and oriented fiber on a mandrel that provides the finished component shape. The mechanics of filament winding are illustrated in Fig. 49.2 The tensioned fiber bands are accurately placed on the mandrel using numerically computer controlled or chain and gear mechanical winding machines. Fig. 49.2 shows the fiber being pulled from tensioned spools through the delivery eye onto the rotating mandrel. The resin can either be pre-applied on the fiber in an off-line operation (pre-impregnation, or wet prepreg) or applied during the winding operation (wet winding). Prepregging usually requires storing the material at low temperature until required for winding to retard the curing process. Wet winding uses a resin
bath, along the fiber path between the tensioning system and the delivery eye, through which the dry fibers are pulled. Both wet winding and prepreg winding are widely used. The choice of technique depends on the application, cost and quality requirements. Composites that have directionally oriented continuous fibers are analyzed by techniques such as netting analysis, rule of mixtures, classical lamination theory, and micromechanics. These analysis techniques all account for the fact that composites have material properties that depend on the fiber directions of each ply. The overall directional properties of the laminate are calculated by adding the relative effects of each ply. The analyses are more complicated than with isotropic materials such as steel and aluminum. However, the widespread use of computers has made the task of analyzing composites relatively simple. Composites are in many cases much stronger and stiffer than metals such as steel, aluminum or titanium. When one takes into account the much lower density of composites as compared to metals, composites have an even greater performance advantage over metals. To compare material properties on an equal weight basis (instead of an equal volume basis), a property such as ultimate strength is divided by the density of the particular material. The resultant is termed specific strength. The same can be applied to stiffness (modulus) and is called specific modulus. Comparing materials this way is very important when designing weight critical parts. Figures 49.3 and 49.4 show the strength, modulus, and specific properties of representative composites and metals. The composite data are for composites with all fibers oriented in one direction (unidirectional ply data). The test loads are applied to the specimens in the same direction as the fibers. Also, all composite data presented here are for 60% fiber and 40% resin by volume samples. This ratio is typical for advanced composite laminates. The epoxy resin used to make the samples is not indicated specifically because its effect on the
Manufacturing techniques 1047
-MACHINE DRIVE
W
orlzontal and VeWtical
Fig. 49.2 Layout of a typical filament winding machine (Courtesyof Composites Machines Company.) unidirectional ply properties is negligible compared to the effect of the fibers. In Fig. 49.3 it can be seen that composite plies are up to twice the strength of high strength steel and three to five times stronger than titanium and aluminum. Comparing specific strengths, composites are three to eight times stronger than the metals. The stiffness of the materials, shown in Fig. 49.4, range from fiberglass composite at the low end to high modulus carbon composite at the high end with metals in between. Composites compare more favorably to metals in specific modulus. The specific modulus of the composites shown are one to six times greater than for metals. Fatigue resistance is another important material property to consider when designing
a product. Fatigue describes the phenomenon of reduced strength over time of load application (static fatigue) or the number of times a load is applied (cyclic fatigue). Aluminum and fiberglass composites suffer from poor fatigue performance whereas carbon compositesexcel. Figure 49.5 displays the cyclic fatigue strength of various metals and composites. Specimens of each material were loaded repeatedly in tension to various predetermined loads and then the loads were removed, until the specimens failed. A curve of cyclic stress against number of cycles to failure can be constructed for each material, from which the 10 million cycle failure stress for each material can be obtained. The 10 million cycle life is considered by designers to approximate an infinite life for most applications.
1048 Composites in the sporting goods industry Material
0
2
Specific Strength, E6 in. 4
8
6
34 msi carbon 42 msi carbon
50 msi carbon 5-2 glass E glass Kevlar Spectra 1000 Titanium 4340 Steel Aluminum
-
0
100
200 Strength, ksi
300
Strength
500
400 Specific Strength
Fig. 49.3 Strength of composites; comparison with metals at 60% fiber volume. Specific Modulus, E6 in.
800
Material 34 msi carbon 42 msi carbon
50 msi carbon S-2 glass E glass Kevlar Spectra 1000 Titanium 4340 Steel Aluminum 0
5
10
15 20 Modulus, Msi
25
-Modulus Fig. 49.4 Stiffness of composites; comparison with metals at 60% fiber volume.
30
35
Specific Modulus
Sporting goods applications 1049 Tension - tension ratio at 10 E + 7 cycles
Material 34 msi carbon 42 msi carbon 50 msi carbon
S-2 glass E glass
Kevlar Titanium 4340 Steel Aluminum
100
200
300 400 500 Alt. stress density, ksi/lb. per cu. in.
600
Fig. 49.5 Fatigue strength of composites; comparison with metals at 60% fiber volume.
The failure loads shown in Fig. 49.5 are because of the changes in 'feel' of the rod. given as stress density which is the test stress Many golf shafts incorporate both divided by the material density. For the mate- carbon/graphite and boron reinforcements to rials tested, composites have two to six times change flex point, reduce shaft weight, increase the specific fatigue resistance as metals. torque resistance and increase clubhead It is important to mention that the previous speed5. comparisons of metals and composites do not tell the whole story. Often, the benefits of a 49.3 SPORTING GOODS APPLICATIONS composite part compared to a metal one would not be as great as indicated by the dif49.3.1 RACKETS ferences in ply strength and stiffness. The design allowable composite ply strength is There are approximately 22.5 million tennis 10-20% lower than the test specimen players and 7 million racquetball players in strengths. Other factors including voids, long- the USA. These players purchased 2.45 milterm environmental effects, use temperature lion tennis rackets at a cost of $170 million and and imperfect fiber alignment all reduce the 1 million racquetball rackets at a cost of $53.4 ultimate strength of composites. Finally, when million retail dollars in 1992. Approximately all the reinforcing fibers are not in the direc- 73% of these rackets were composite4. Most all of these rackets are produced in tion of the load, the fibers do not carry the load as efficiently. However, when these factors are the Far East using a compression molding adequately controlled, there are many applica- process with an internal bladder. This process tions where composite structures greatly consists of hand lay-up of prepreg around a outperform their metal counterparts. Sport straight metal mandrel and plastic bladder. and recreation applications are one market After lay-up the prepreg/bladder is removed area where composites have a decided advan- from the mandrel and placed in a compression mold. The mold is closed, the bladder is tage over metals. There are other not easily quantifiable rea- pressurized, and the part is temperature sons for choosing a particular composite or an cured. A significant portion of the labor in individual lay-up. Fishmg rods generally do manufacturing a racket is in the surface finnot have hoop reinforcement or angle plies ishing, painting and labeling.
1050 Composites in the sporting goods industy Other processes that have been used include injection molding, braiding, and resin transfer molding and combinations of these processes. Filament winding is being considered because it can allow the use of wet winding, which reduces the raw materials costs and labor content in racket manufacturing. This process uses wet resin and dry fibers and provides product consistency because the operation is machine controlled. The preform can be wound on a straight mandrel to allow standard filament winding techniques to be used. After winding, the preform can be removed from the mandrel and placed in the mold. The key technology is to place the wet fiber preform in the mandrel without disturbing or kinking the fibers. The process then continues as the typical processes now being used. A typical product of this method is shown in Fig. 49.6. Filament winding can also be used to dry wind the fiber to make a cheap oriented fiber preform. Then using RTM, the resin can be injected with the prospect of further reducing hand surface finishing. An additional advantage to filament winding is the design flexibility of changing strength and stiffness around the Fig. 49.6 Prince graphite extender racket. (Courtesy racket head, which is difficult to accomplish of Spyrotech.) using lay-up techniques. impact on the acceptance of carbon fiber composites in the bicycle industry. In 1991 more 49.3.2 BICYCLE FRAMES AND COMPONENTS than 50% of the carbon bicycles produced Advanced composite bicycle frames first worldwide (>lo0 000) were manufactured in appeared in the early 1970s. These frames Taiwan6. Tube and lug designs dominated the marwere not successful due to poor design and manufacturing techniques and market resis- ket until 1993. The technique consists of tance to the new materials. In the mid-l980s, bonding premade tubes into forged or cast composites in bicycles had success in the mar- metal lugs. Trek developed a composite lug ket place and today are the materials of choice that is offered in some of their frames. Kestrel for new designs6. For the past few years introduced a molded monocoque frame in approximately 12 million bicycles were sold 1987. Molded monocoques are manufactured annually in the USA. Of these, approximately by lay-up of composite material into female 10% were considered high-end (sales price mold halves with the reinforcing metal pieces as required. The female mold sections are over $750) products. In 1989 and 1990 Greg LeMond won the brought together and internal bladders are Tour de France Bicycle Race using a carbon inflated to provide compaction pressure durfiber frame. His success has had a significant ing the curing operation.
Sporting goods applications 1051 Tube and lug frames are relatively simple to manufacture. However, the frames can have durability problems if the tubes and tube to lug joint are not properly designed to prevent significant stiffness mismatches. Galvanic corrosion potential is a concern. Filament winding offers an inexpensive method for manufacturing frame tubes. Wet winding techniques allow using the required materials in the lowest cost form. Since the process is machine controlled, labor requirements are minimized. Currently the most widespread manufacturing method is roll-wrapping. Rollwrapping requires cutting prepreg plies of composite and hand wrapping them around a metal mandrel prior to curing. Frame tubes are not the only use of composites in the bicycle industry. Handlebars are filament wound, removed from the mandrel, and then cured in a mold to provide the proper shape and curvature. An internal bladder provides the compaction pressure. Seat posts, swing arms, and handlebar extensions can be made using this process. Titanium tubes in some front wheel suspension systems are reinforced with composite tubing. Wheels and spokes are also being made of composite materials using combinations of roll wrapping, hand lay-up and resin transfer molding.
helps to reduce the seam effect on performance consistency. Grafalloy introduced a filament wound shaft for the 1993 season. Filament wound shafts provide superior performance at low weight and cost compared to roll-wrapped shafts available with comparable performance specifications. Those familiar with the industry believe filament winding will be the process of choice in manufacturing golf shafts in the next few years. Player tests and mechanical golfer testing have proven the performance improvements of a filament wound shaft over roll-wrapped shafts. Many original equipment manufacturers are evaluating filament wound shafts for their product lines. One factor where rollwrapped shafts can have an advantage over filament wound shafts is in tip durability. Filament winding can introduce more voids in the composite laminate than roll-wrapping. Special attention is needed by filament winders to insure that their product has sufficient tip strength. ASTM is working to establish test methods for golf shafts. Standard tests will allow for a better comparison of products and give the customer more confidence in the quality and performance of the product being purchased.
49.3.3 GOLF SHAFTS
49.3.4 FISHING POLES
In 1992 over 100 million composite golf shafts were made world wide with a wholesale dollar value of $900 million7.The majority of the shafts were manufactured in the Far East and sold for less than $10 each wholesale. At the present time, virtually all composite golf shafts are manufactured using the roll-wrap process. This process relies on manual orientation of the plies of material on the mandrel and ply alignment can be inconsistent from part to part. The alignment directly affects the shaft performance; flex, and torque. The cut plies also create a seam down the length of the shaft which causes inconsistency in the shaft performance. Attention to ply alignment
In 1992, 16 million fishing poles were purchased in the USA. Virtually all were produced in the Far East using the roll-wrap process. It is estimated that about 20% are high-end products and may be amenable to the manufacturing costs of automated processes in the USA. For filament winding, this market segment is the least developed. Development work is underway to produce a cost competitive high-end fishing pole in the USA. 49.3.5 OTHER APPLICATIONS
Other applications for composites in the sporting goods industry include softball bats, pool
1052 Composites in the sporting goods industry cues, kite tubing, shoe inserts, ski poles and model aircraft. The majority of these products are currently being manufactured by either roll-wrapping, hand lay-up, or pultrusion. Improvements in material properties, reduction in material prices, and significant improvement in component performance has created a market for composites in the sporting goods industry*. Unfortunately for USA manufacturers, the majority of the sporting goods products are fabricated overseas. This business will only come back to the USA if companies are cost competitive and/or offer a product of higher value which is superior in performance. One method to improve cost competitiveness is through the use of automation. Filament winding, along with innovative engineering, offers the possibility of bringing composites to the forefront in the sporting goods industry.
REFERENCES
1. Segal, C.L. Worldwide Markets For Advanced Fibers, 23rd Intern. SAMPE Tech. Conf., Kiamesha Lake, NY, October 23,1991. 2. Anon. Westinghouse ESG News, Oct, 1994. 3. Kline & Company The Changing Advanced Polymer Composites Market, 1993. 4. Feeney, B. Composites in Racquetball and Tennis Rackets, SME Conf. Composites Manufacturing and Tooling '94, Anaheim, California, January 17-20. 5 . Levin, S. The Use of Composite Materials in the Bicycle Industry, SME-Effective Applications of Composites in the Sporting Goods and Recreational Industry Clinic, August, 1992. 6. Textron Specialty Materials, Boron Backgrounder. 7. McConnell, V. Composites in sports: goingfor the gold, Adv. Comp., September/October 1992. 8. Mcconnell, V. Sports Applications - Composites at Play, High Performance Comp., Jan/Feb 1994.
TYPICAL PROPERTIES FOR ADVANCED COMPOSITES
APPENDIX A
Kenneth R. Berg
A.l INTRODUCTION
material and configuration is selected, a minimum test program would then be initiated. For a company or institution that is designing Having a set of typical composite materials composite material structures, or embarking has advantages and disadvantages. For examfor the first time into the application of ple, if one were to design a structure utilizing advanced composite materials for structural only typical material properties, without the purposes, it is imperative that material propknowledge of the scatter that may occur in erties be available. Of course it would be those properties, structural failure may occur. desirable to have a complete set of statistical Perhaps not immediately, nor on every strucDesign Allowables, such as the statistical 'A' ture produced, but on an unknown statistical values for properties, or even the 'B' values, basis, at some point in time. However, prior to (see Chapter 33 for detailed definitions of a final design for a structure, the normal engithese values and Neal and Spiridgliozzi, 1987). neering procedure is to initiate the test Since complete statistical Design Allowables program. The purpose of the test program is are not available, the next sought after materthreefold: one, confirmation of the design; ial properties would be 'typical' properties. two, determine the scatter that occurs due to However 'typical' properties are not defined variations in materials and the manufacturing statistically and may be defined in many difprocess; and three, over a period of time, ferent ways. Therefore it is important to either to confirm the material properties datadiscuss typical material properties and also base being used, or to accumulate test data for discuss the means to achieve a set of typical a material properties database. proper ties. The purpose of having a complete set of typical properties is to be able to design com- A.2 TYPICAL PROPERTIES - CONSTITUENTS posite structures with a minimum of testing confirmation. Having a complete set of typical A.2.1 FIBERS properties will allow design optimization, preliminary design, cost and weight optimization One of the problems of determining typical and other trade-offs with a number of different properties is the variations that occur in the materials and candidate laminates with differ- materials making up laminates. In the case of ent fiber orientations. Once an optimum glass fiber, the types of glass fiber and number of manufacturers is considerably less than with carbon fiber. However, even with this limitation, there are Handbook of Composites.Edited by S.T. Peters. Published at least two major types of glass fiber, E-glass in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7
1054 Typical properties for advanced composites and S-2 glass. Withm each of these glasses are variations in chemical composition, fiber diameter, fiber finish, fiber sizing, the number of individual fibers in a tow, roving, yarn, etc. Manufacturers have different names for the similar type of glass, for example the higher strength, higher modulus glasses. These fiber glasses are the older S-glass (no longer available commercially), s-2 glass and the R-glass by a French manufacturer. Other countries fabricate the same type of glass, but with only minor differences in properties. For carbon fiber, not only are there the same variations as mentioned above for glass, but in addition, there are large variations in strength and modulus and in manufacturers, see Tables A.l and A.2. Based on the large number of variations in fibers, it would be virtually impossible to obtain complete statistical material properties for each variation. Even to obtain typical properties for each variation would not be practical. To reduce this problem to a practical level, it is necessary to analyze the usage of glass and carbon fibers (or other fibers). The usage of advanced composite fibers by ‘pounds used per dollar expended’, is estimated to be, in order of highest usage: E-glass, high strength carbon (modulus of 227 GPa, (33 x lo6 psi)) and then S-2 glass. With this list, it is possible to develop typical properties for composites fabricated from each of these fiber types. An example of the determination of the strength and modulus of the typical high strength carbon fiber is shown in Fig. A.l. The ’typical’ property becomes: Tensile modulus of 227 GPa (33 x lo6 psi), and tensile strength of 4000 MPa (580 ksi). A.2.2 RESIN SYSTEMS
The matrix for fiber composites can be classified into two categories, metallic and non-metallic. This discussion on typical properties involves only non-metallic resin matrix systems.
Table A.l Typical products from carbon fiber manufacturers (available in USA) A M O C O (Thornel)”
Toho Rayon (Besfight)
P-100 P-75 T-300 T650/35 T650/42
HTAWlOO im600 HM35
Toray (Torayca)
G30-500 G40-800 G40-600
T300 t7005 T800H M40 M46J
Grafil (Grafil)b
AKZO (Fortafil)
34-700 42-650 42-750
F-5 F-3
Toho (Celion)
-
Hexcel (Magnamite) AS4 IM6 IM7 IM8 a
Typical product name Grafil is a subsidiary of Mitsubishi Rayon Co. Ltd and their fibers are called Grafil or Pyrofil.
Table A.2 Torayca fiber types (Toray, 1991) Fiber type
Number offilaments
T300 T300J T400H T700S T800H TlOOOG TlOOO M35J M40J M46J M50J M55J M60J M30 M30SC M40 M46 M50
lK, 3K, 6K, 12K 3K, 6K, 12K 3K, 6K 12K 6K, 12K 12K 12K 6K, 12K 6K, 12K 6K, 12K 6K 6K 3K, 6K lK, 3K, 6K, 12K 18K lK, 3K, 6K, 12K 6K lK, 3K
Typical properties - constituents 1055 300
Typical Property
Typic.i
Prop.rt1.r
Strength 4 GPa Modulu6 22.0 GPO
n 20 a
Modulus 227 GPa
0
g 100 Data horn various Rbsr manufacturers
5 (I 0 0 0
+ EPoxy/Anhydrld*
' EpoxyIAmlne
5
Polymstw
1
2
3
4
5
0
100
ZOO
Vlnyl*stwr
bpisal Property
300
400
500
800
700
800
900
1,000
Strength MPa
Strength GPa
Fig. A.l Typical strength and modulus for high
Fig. A.2 Typical strength and modulus for E-glass
strength carbon fiber. (Courtesy of Eggs Corporation.)
composite - flexural strength. (Courtesy of Riggs Corporation.)
As was discussed for fibers, only the high usage matrix systems in advanced composites are considered as candidates for typical properties. In addition, for typical properties of advanced composites for structural applications, only structural resin systems are candidates. Structural resins are defined as resins that have similar modulus and tensile strength as standard epoxy systems. For example, an applicable resin for structural composites would have a modulus of approximately 3.5GPa (0.5 x 106psi) and a tensile strength of approximately 100 MPa (15 h i ) . The more popular structural resins are polyester, epoxy, vinyl ester and phenolic. For typical composite properties, the use of any of these resins will allow a single typical property (Fig. A.2) (CertainTeed Corporation, 1989). A comparison for composites with different epoxy sizing from different manufacturers and a typical value is shown in Fig. A3. Figure A.4 shows a number of different resin systems and the typical values for strength and modulus (Lubin, 1987).This data is for primarily fiber controlled properties. There are properties in which the resin is the significant factor. These characteristics are associated with stress concentrations and
environmental considerations. Key among these characteristics are: temperature, fracture toughness, compression after impact, crack propagation, humidity, stress concentrations, interlaminar shear, mechanical fasteners in laminates, holes in laminates, creep, damage tolerance and compatibility with fiber finish. In determining typical properties, these characteristics are not included but, as applicable, need to be considered for the final design.
cp o o / L
-
400
0
t
300
* Cellon G30/6-F Herculee AS-4mpe 0
X
1
Hercules AS-4ffreated
0
Fig. A.3 Typical compression strength for carbon fiber fabric composites. (Courtesy of Riggs
Corporation.)
1056 Typical properties for advanced composites 100
90 80
2
70
-2
60 50 40
B 30 20 10
-
0
0
100
200
300
400
500
600
700
Tensile Strength MPa Fig. A.4 Typical strength and modulus for carbon fiber fabric composites. (Courtesy of Riggs Corporation.) A.3 TYPICAL PROPERTIES - COMPOSITES
A.3.1 FIBER CONTROLLED TYPICAL PROPERTIES
Fiber-reinforced composite materials are primarily used to take advantage of the high strength and stiffness of the fiber. Therefore in most applications, the laminate orientation is designed so that the strength and modulus are controlled by the fiber properties. For example, for a typical fiber orientation in a laminate of 0"/&"/90", the 0" plies control the failure of the laminate whenever the percentage of 0" plies is greater than lo%, (with a I3 greater than *loo). For I3 less than do", if the combined percentage of 0" and &"plies is greater than lo%, then the laminate is also controlled by fiber fracture. These limitations are approximate and depend on the fiber strength and modulus, but are valid for carbon, aramid fibers and glass. For high I3 angles and for 90" laminates, the tensile strength of the laminate is reduced below the tensile strength of the resin due to stress concentrations between the fibers (Berg, 1967). Figure A.5 shows the tensile strength of a carbon fiber laminate of 0"/+eo/9O0 plies, with the strength of the resin varying to influ-
ence the transverse strength and modulus of the base unidirectional laminate (100% 0" plies). The different strength resins shown on Fig. A.5, are as follows: Transverse modulus (GPa) Lower strength resin Typical strength resin Higher strength resin
700
Transverse tensile strength (MPaa) 41.4 55.2 69.0
8.6 10.0 11.4
Laminate 15% 0",15% f Oo, 70% 9 0
I
0
30
15
46
Angle 20'
Fig. A.5 Effect of resin properties on tensile strength of high strength carbon composites. (Courtesy of Riggs Corporation.)
Selecting the use of typical properties As can be seen in Fig. A.5, the typical fiber A.4 SELECTING THE USE OF TYPICAL fracture composite properties (ultimate fail- PROPERTIES ure) are not affected by large variations in the properties of the resin. However, for resin sen- A.4.1 WHERE TYPICAL PROPERTIESARE sitive properties, for example, the first ply ACCEPTABLE failure (limit) of the 90" plies, the resin propThe acceptability of material properties in the erty is influential. design of structures is based on a number of For the fiber fracture controlled composites, factors. If a design is being produced for a cuswhich are the main interest in structures, typitomer, the customer is often the final word on cal composite material properties are valid the acceptability of the material properties utiover a wide variation in resin characteristics. lized. An alternate possibility is that the customer is not interested in accepting or A.3.2 MATRIX CONTROLLED TYPICAL rejecting the material properties used, but PROPERTIES would rather accept the product against a The main interest in structural components is specification. Final acceptance is a qualificathe high strength and modulus obtainable tion of the product through testing. from the fiber reinforcing of the matrix. The third case is where a product is proTherefore in the design of the laminate, for duced by the company itself and sold to the most of the applications, the resulting failure consumer directly. Of course the consumer modes are fiber fracture critical. Unfortunately, (public) is not interested in accepting or rejecthere are cases where the matrix is the critical tion the material properties database. An element in the failure mode for all laminate example of this is the automobile industry. orientations. This does not mean that the fiber In the case of a company that subcontracts does not contribute to reinforcing the matrix in the design and fabrication of composite strucboth strength and modulus, but only that the tures, the company may either want to review ultimate failure is the result of failure in the the subcontractor's material properties datamatrix. base, or be able to review the design of the Among the cases where the matrix is the subcontractor with the use of the company's critical failure mode are laminates that are typical properties database. subjected to shear, the first ply failure (limit In each of the above cases, a typical materload) of a laminate (Fig. A.5, first ply failure ial properties database can be used for cost curves and Berg, 1982) and most cases of the and weight trade-offs, selection of the best transverse strength property of a laminate materials, optimizations studies and prelimiwith no 90" plies. Even for these cases, a typ- nary design. ical property can be determined utilizing the It is important to note that the final design typical set of unidirectional properties repre- would always go through an extensive test sented by most of the epoxy systems in use program to verify the material properties by the prepreg manufacturers as well as most selected, the final design, the manufacturing of the structural epoxies sold. A typical set of process and to determine the variability of the unidirectional properties for high strength product. carbon composites, for example would be: Thus, a typical material properties database is acceptable and useful to reduce the cost of 0" 90" Strength engineering design, reduce the cost of testing Tension 2100 MPa 55 MPa Compression 1380 MPa 190 MPa and allow a more intelligent and less time con55 MPa In-plane shear suming approach to the final design. Modulus Shear modulus
9.7 GPa
138 GPa
6 GPa
1058 Typical properties for advanced composites A.4.2 WHERE TYPICAL PROPERTIES ARE NOT ACCEPTABLE
the composite transformations equations (Tsai, 1988). A typical property database for all orientations and selected materials is available In the above discussion, the customer require(Berg, 1993), but also a typical composite matements were mentioned as one of the criteria rial database could be generated by a for acceptance or rejection of a material propcompany using composite engineering analyerties database. There are cases where the sis as discussed above. customer will insist upon enough testing to Although this discussion of typical properdevelop a statistical property database. For ties has mentioned only static strength and that requirement the typical material database modulus, typical properties are also available would not be acceptable. However in any stafor fatigue, CTE and for elevated temperatures tistical database, the applicability is confined (Berg, 1993). Fatigue typical properties include to the specific fiber, matrix and fabrication all fatigue stress ratios. Fatigue statistical process. The statistical testing process is properties, of course, would be prohibitively expensive and time consuming. In most cases, expensive. the statistical database would be limited to the If users of typical composite material proplaminate orientations tested, which would also erties are aware of their limitations, typical be very limited. properties can be a very useful database for For applications where sigruficant environcost effective design and analysis. mental conditions are present, the use of typical material properties may not be applicable due to large variations in the response of different REFERENCES resins to these environmental conditions. These conditions were summarized above and Berg, K.R. 1967. The effect of fiber spacing on the strength of composites laminates, Paper preincluded impact, humidity other corrosive flusented at AAIAIASME 8th Structures, Structural ids, stress concentrations, temperature, etc. Dynamics and Materials Conference, Palm Springs, CA. A.5 SUMMARY Berg, K.R. 1982. The effect of matrix strain limitations on composite design allowables. Proc. 27th The design of composite material structures Nat. S A M P E Symp., May, 1992, San Diego, CA. requires a knowledge of the material properBerg, K.R. 1993. Composite Material Properties Data ties for all combinations of laminates. It is Books - Sample Data Sheets. RIGGS Corporation, cost-prohibitive to test all combinations of Seattle, WA. laminates, even to obtain typical properties. To CertainTeed Corporation. 1989.Sales Brochure, Test Data, Wichita Falls, TX. obtain statistical design properties for a limited number of laminate configurations is also Hashin, Z. and Rotem. A. 1975. Failm modes of angle ply laminates. I. CompositeMater., Volume 9. expensive, but in some cases may be required Lubin, G. (ed.) 1987. Handbook of Composites. New by a contract. Typical composite material York: Van Nostrand Reinhold. properties can provide useful data and be cost Neal, D. and Spiridgliozzi, L. 1987. An efficient effective for the design engineer. The data can method for determining the 'A' and 'B' design be generated by utilizing typical unidirecallowables. Army Materials and Mechanics tional data for each class of materials, (as Research Center ,Watertown, MA. discussed above for high strength carbon com- Toray Industries, Inc. 1991. Toruycu Sales Brochure, Typical Fiber Properties. posites). To generate all the laminate Tsai, Stephen W. 1988. Composites Design. Dayton, configuration, both limit (first ply failure) and OH: Think Composites. ultimate, requires a comprehensive computer Tsai, Stephen W. and Hahn. H. 1980. Introduction to program, including a failure criteria (Tsai, Composite Materials. Westport, CT Technomic 1988; Hashin and Rotem, 1975) and utilizing Publication.
SPECIFICATIONS AND STANDARDS FOR POLYMER COMPOSITES
APPENDIX B
Frank T. Traceski
B.l USES OF SPECIFICATIONS AND STANDARDS
Material specifications and engineering standards for advanced composite materials have a very broad applicability across the entire spectrum of defense and commercial applications. From basic research, through engineering and manufacturing development, in production, and for maintenance and field repair, material and process specifications establish requirements and procedures. In research and development, testing standards are used for material characterization to determine physical, chemical, mechanical, thermal and electrical properties. In manufacturing development, material specifications are used to establish material quality and processability requirements. In production, the material producer uses test standards and material specifications for statistical process control (SPC)to ensure batch-to-batch consistency. Material testing standards are used extensively in design engineering to determine material design allowables (i.e. strength and stiffness limits beyond which catastrophic failure occurs) at various temperatures and environments. Material testing to standards enables one to quantify material performance Handbook of Composites.Edited by S.T. Peters. Published in 1998 by Chapman &Hall, London. ISBN 0 412 54020 7
limits. A list of typical composite material qualification tests is provided in Table B.l. These examples are representative of the material properties which are determined in a typical material qualification program in accordance with various test standards. A concurrent engineering approach to composites engineering requires that material selection (as part of the design process) be integrally linked with engineering and manufacturing process development. In composite manufacturing development, material process specifications are defined for a given composite material and manufacturingprocess. Once optimized, process specifications reduce manufacturing risk and enhance producibility. Another aspect of composites manufacturing is the nondestructive testing and inspection (NDTI) of composite structures to verify structural integrity during production. Ultrasonic inspection, infrared thermography, and a host of other NDTI methods are employed to locate voids, delaminations, cracks, and other types of structural defects. Standards are also employed here to define NDTI procedures and acceptance criteria. In summary, engineering test standards and material and process specifications are employed extensively in composites research, development and production. Appendix B identifies specific material specifications and test standards for polymer composites and the
1060 Specifications and standards for polymer composites Table B.1 Typical composite material qualificationtests
Physical
Mechanical
Chemical
Resin content Resin areal weight Resin flow Glass transition temperature
Tensile strength and modulus Compression strength and modulus Short beam shear (SBS) 245 In-plane shear Open-hole tension Open-hole compression Compression after impact (CAI) Flexural strength and modulus Fatigue strength Creep Dynamic mechanical analysis (DMA) Instrumented impact (toughened resins) Fracture toughness (toughened resins) Solvent sensitivity compression strength, SBS Bolt bearing
Infrared spectroscopy (IR) Liquid chromatography (HPLC) Hydraulic fluid/solvent resistance Fuel (JP-4)resistance
(Tg)
Rheological dynamic spectroscopy (RDS) Gel time Volatiles content Fiber density Fiber mass per unit length Fiber content Prepreg tack Laminate ply thickness Laminate void content Laminate density Laminate fiber volume Laminate flammability
Thermal Thermogravimetricanalysis (TGA) Thermomechanical analysis (W) Differential scanning calorimetry (DSC) Thermal oxidative resistance Thermal expansion (CTE) Thermal cycling Thermal conductivity Electrical Dielectric constant Dielectric strength Dissipation factor Volume resistivity Dielectrometry
Some of these tests are specific to resin, fiber reinforcement, prepreg or laminate. There is no established universal set of qualification test procedures which is widely adopted. See MIL-HDBK-17 for recommended guidelines.
organizations that develop them. It is limited to consensus-type standards and does not include private sector specifications which are generally not available. B.2 STANDARDS-DEVELOPING ORGANIZATIONS
The two principal organizations which develop test methods for composites in the USA are the American Society for Testing and Materials (ASTM) and the Suppliers of Advanced Composite Materials Association (SACMA). The Society of Automotive Engineers (SAE) Polymeric Materials Committee is the organization which publishes Aerospace Material Specifications (AMS) for advanced polymer composites. The Department of Defense (DoD) has also issued
several military specifications and standards for polymer composite materials. Other private sector organizations, such as the Aerospace Industries Association (AIA) and Composite Materials Characterization (CMC), Inc., are involved in the standardization of composite materials and tests in order to reap long-term economic savings. Japan, Germany, France and UK are also major players in composites technology. The European Association of Aerospace Manufacturers (known as AECMA in Europe) produces European Norm (EN) standards for aerospace composites. Japanese Industrial Standards (JIS) serve as a basis for standardization of composites in Japan. Germany has issued DIN standards for composite materials. France has AFNOR standards and the UK has British Standards.
Standards used in USA 1061 Engineering standards for polymer composites also promote international commerce. In the global arena, the International Standardization Organization (ISO) is the body which develops international standards. The ISO/TC 61 Subcommittee (SC) 13 on Composites and Reinforcements Fibers is currently coordinating new standards for glass and carbon fiber composites. USA participation with IS0 not only promotes international commerce but also enhances USA global competitiveness in the composites industry. B.3 STANDARDS USED IN USA
B.3.1 ASTM STANDARDS
The ASTM Committee D30 on High-Modulus Fibers and their Composites develops standard test methods for advanced polymer composites. Table B.2 lists the principal ASTM
standards for determining the physical and mechanical properties of polymer composites. ASTM standards are developed by a consensus process and are widely used. The ASTM Committee D20 on Plastics also has developed standards which may be used for testing plastic resins and reinforced plastics. 8.3.2 SACMAMETHODS
The SACMA has developed recommended test methods for determining the physical, mechanical, and chemical properties of composite materials (Table B.3). Although SACMA is not a standards-setting body, it works actively with ASTM, SAE, ISO, DoD, AIA and others towards standardization of composite test methods. SACMA does not regard its SRMs to be ’standards’ in the truest sense because they have not been developed by a consensus process which is typical of standard-developing bodies.
Table B.2 ASTM Standards for advanced polymer composites ASTM C613 ASTM D695 ASTM D790 ASTM D2290 ASTM D2344 ASTM D2734 ASTM D3039 ASTM D3171 ASTM D3379 ASTM D3410 ASTM D3479 ASTM D3518 ASTM D3529 ASTM D3530 ASTM D3531 ASTM D3532 ASTM D3544 ASTM D3800 ASTM D3878 ASTM D4018 ASTM D4102 ASTM D4255 ASTM D5229 ASTM D5300 ASTM D.5467 ASTM D5528
Resin Content of Carbon and Graphite Prepregs by Solvent Extraction Compressive Properties of Rigid Plastics Flexural Properties of Unreinforced and Reinforced Plastics Plastics, Ring or Tubular, Apparent Tensile Strength of, By Split Disk Method Apparent Interlaminar Shear Strength of Parallel Fiber Composites by Short-Beam Method Void Content of Reinforced Plastics Tensile Properties of Polymer Matrix Composite Materials Fiber Content of Resin-Matrix Composites by Matrix Digestion Tensile Strength and Young’s Modulus for High-Modulus Single-Filament Materials Compressive Properties of Polymer Matrix Composite Materials with Unsupported Gage Section by Shear Loading Tension-Tension Fatigue of Oriented Fiber, Resin Matrix Composites In-Plane Shear Response of Polymer Matrix Composite Materials by Tensile Test of a +/- 45 D e p e Laminate Resin Solids Content of Epoxy-Matrix Prepreg by Matrix Dissolution Volatiles Content of Epoxy Matrix Prepreg Resin Flow of Carbon Fiber-Epoxy Prepreg Gel Time of Carbon Fiber-Epoxy Prepreg Reporting Test Methods and Results on High Modulus Fibers Density of High-Modulus Fibers Standard Terminology of High-Modulus Reinforcing Fibers and their Composites Properties of Continuous Filament Carbon and Graphite Tows Thermal Oxidative Resistance of Carbon Fibers In-plane Shear Properties of Composite Laminates Moisture Absorption Properties and Equilibrium Conditioning of Polymer Matrix Composite Materials Measurement of Resin Content and other Related Properties of Polymer Matrix Thermoset Prepreg by Combined Mechanical and Ultrasonic Methods Compressive Properties of Unidirectional Polymer Matrix Composites Using a Sandwich Beam Mode I Interlaminar Fracture Toughness of Unidirectional Fiber-Reinforced Polymer Matrix Composites
1062 Specifications and standards for polymer composites Table B.3 SACMA Recommended methods (SRM) Compressive Properties of Oriented Fiber-Resin Composites Compression After Impact Properties of Oriented Fiber-Resin Composites Open-Hole Compression Properties of Oriented Fiber-Resin Composites Tensile Properties of Oriented Fiber-Resin Composites Open-Hole Tensile Properties of Oriented Fiber-Resin Composites Compressive Properties of Oriented Cross-Plied Fiber-Resin Composites In-plane Shear Stress-Strain Properties of Oriented Fiber-Resin Composites Short Beam Shear Strength of Oriented Fiber-Resin Composites Tensile Properties of Oriented Cross-Plied Fiber-Resin Composites Fiber Volume, Percent Resin Volume and Calculated Average Cured Ply Thickness of Plied Laminates Environmental Conditioning of Composite Test Laminates Lot Acceptance of Carbon Fibers Printing and Applying Bar Code Labels Determination of Mass Per Unit Length of Carbon Fibers Determination of Sizing Content on Carbon Fibers Determination of Density of Carbon Fibers Tow Tensile Testing of Carbon Fibers Determination of Twist in Carbon Fibers Glass Transition Temperature Determination by DMA of Oriented Fiber-Resin Composites Viscosity Characteristics of Matrix Resins High Performance Liquid Chromatography of Thermoset Resins Fluid Resistance Evaluation of Composite Materials Determining the Resin Flow of Preimpregnated B-Staged Materials Determination of Resin Content and Fiber Areal Weight of Thermoset Prepreg with Destructive Techniques Determination of Resin Content, Fiber Areal Weight and Flow of Thermoset Prepreg by Combined Mechanical and Ultrasonic Methods SRM 25 Heat of Reaction, Onset Temperature and Peak Temperature for Composite System Resins Using Differential Scanning Calorimetry (DSC) SRM 26 Fiber/Matrix Adhesion of Carbon Fiber Reinforced Polymer Matrix Composites
SRM 1 SRM 2 SRM 3 SRM 4 SRM 5 SRM 6 SRM 7 SRM 8 SRM 9 SRM 10 SRM 11 SRM 12 SRP 1 SRM 13 SRM 14 SRM 15 SRM 16 SRM 17 SRM 18 SRM 19 SRM 20 SRM 21 SRM 22 SRM 23 SRM 24
Table B.5 lists military specifications for various fiber reinforcements and composite materials. The Military Handbook 17 effort is the most widely r e c o h e d DoD standardization project. The purpose of this handbook is to provide B.3.3 AEROSPACE MATERIAL SPECIFICATIONS a SOUTCe of based
However, SACMA recommended methods (SRMs) are being used as standards by various organizations. -
The Society of Automotive Engineers (SAE) is the primary organization in the USA which issues material specifications for polymer composites. Table B.4 lists some typical Aerospace Material Specificationsfor -polymer composites. In general, the SAE has published specifications for carbon, aramid, glass and boron fiber composites. B.3.4 MILITARY SPECIFICATIONS
The Department of Defense (DoD) has issued several military specifications for polymer composite materials used in aircraft applications.
ical property data for current and emerging composite materials. The MIL-HDBK-17 government/industry coordination group meets twice annually to develop this handbook. B,3.5 NASA STANDARDS
The National Aeronautics and Space Administration (NASA) has developed five standard tests and a material specification for carbon (graphite) composites. Table B.6 lists the specific tests and the NASA/aircraft industry specification for toughened epoxy composite materials.
Standards used in USA 1063 Table 84 Aerospace material specifications (AMS) for polymer composites
CarbonEpoxy Composites AMS 3892B Fibers, Carbon Tow and Yam, for Structural Composites AMs 3894E Carbon Fiber Tape and Sheet, Epoxy Resin Impregnated AMS 3895B Broadgoods and Tape, Multi-Ply Carbon Fiber/Epoxy, Resin Impregnated, Uniform Fiber
Aramid/Epoxy Composites AMS 39018 Organic Fiber (Para-Aramid),Yarn and Roving, High Modulus AMS 3902C Cloth, Organic Fiber (Para-Aramid),High Modulus for Structural Composites AMS 3903A Cloth, Organic Fiber (Para-Aramid),High Modulus, Epoxy Resin Impregnated
GlassEpoxy Composites
-
AMS 3821B Cloth, Type 'E' Glass, ' B Stage Epoxy-Resin-Impregnated,7781 Style Fabric, Flame Resistant AMS 3828C Glass Roving, Epoxy-Resin-Impregnated,Type 'E' Glass AMs 3831A Cloth, Type 'E' Glass, ' B Stage Epoxy Resin Impregnated, 7781 Style Fabric, Flame Resistant, Improved Strength AMS 3832C Roving, Type '$2' Glass, Epoxy Resin Impregnated AMs 3906B Glass Tape and Flat Sheet, Non-Woven Cloth, Epoxy Resin Impregnated, For Hand and Machine Layup
Boron/Epoxy Composites
AMs 3865C Filaments, Boron, Tungsten Substrate, Continuous AMs 3867I3 Boron Filament Tape, Epoxy-Resin-Impregnated
Division is the Materials and Structures Committee (AIA/MSC) which is responsible The Federal Aviation Administration (FAA) for the coordination and review of proposed has issued Advisory Circulars (AC) to assist specification requirements for materials, commercial aircraft manufacturers in demonprocesses and structures. To promote stanstrating compliance with the requirements of dardization AIA has initiated Project 340-1 the Federal Aviation Regulations in the design Standardization of Advanced Composite and manufacture of composite material strucMaterials and has issued two National tures. Table B.7 lists two Advisory Circulars Aerospace Standards (NAS) for composites issued by the FAA. As of this writing, AC manufacturing which are listed in Table B.8. 145-6 is a draft document undergoing coordination. AC 145-6 addresses requirements for composite repairs, including materials, 8.3.8 COMPOSITE MATERIALS CHARACTERIZATION,INC. (CMC) processes, and quality control tests. B.3.6 FAA ADVISORY CIRCULARS (AC)
Composite Materials Characterization, Inc. (CMC) is a joint enterprise funded by users of B.3.7 AEROSPACE INDUSTRIES ASSOCIATION advanced aerospace composite materials. CMC (AIAI was formed as a result of an Aerospace The Aerospace Industries Association (AIA) Industries Association (AM) initiative to proplays a lead role in composites standardiza- mote industry research collaboration.CMC is a tion. Within the AIA Engineering Standards Delaware corporation chartered in 1987 to
1064 Specifications and standards for polymer composites Table 8.5 Military specifications and standards for polymer composite materials Yam, Cord, Sleeving, Cloth, and Tape-Glass Core Material, Plastic Honeycomb, Laminated Glass Fabric Base, For Aircraft Structural and Electronic Applications Sandwich Construction, Plastic Resin, Glass Fabric Base, Laminated Facings and Honeycomb MIL-S-9041B Core for Aircraft Structural and Electronic Applications Cloth, Glass, Finished, For Resin Laminates MIL-C-9084C Plastic Laminate and Sandwich Construction Parts and Assembly, Aircraft Structural, Process MIL-P-94OOC Specification Requirements Plastic Laminates, Fibrous Glass Reinforced, Marine Structural MIL-P-17549D Cloth, Woven Roving, For Plastic Laminates MIL-C-19663D Thermosetting Polymer Composite, Unidirectional Carbon Fiber Reinforced Prepreg Tape MIL-T-29586 (AS) (Widths Up to 60 Inches), General SpecificationFor Mat, Reinforcing Glass Fiber MIL-M-43248C Armor, Lightweight, Ceramic-Faced Composite MIL-A-46103D Armor, Woven Glass Roving Fabrics MIL-A-46165 (MR) Plastic Laminates, Glass Reinforced (For Use in Armor Composites) MIL-I-46166 (MR) Plastic, Sheet Molding Compound, Polyester, Glass Fiber Reinforced (For General Purpose MIL-P-46169A Applications) Prepreg, Unidirectional Tape, Carbon (Graphite) Fiber Polyimide (PMR-15) Resin MIL-P-46187 Impregnated, 316 C (600 F) Prepreg, Woven Fabric, Carbon Fiber Bismaleimide (BMI) Resin Impregnated MIL-P-46190 Laminate: High-Strength Glass, Fabric-Reinforced, Polyester Resin Preimpregnated MIL-PRF-46197A Roving, Glass, Fibrous (For Prepreg Tape & Roving, Filament Winding & Pultrusion MIL-R-60346C Applications) Armor: Aluminum-Aramid, Laminate Composite MIL-A-62473B Fabric, Carbon (Graphite) Fiber, Nickel-Coated MIL-F-64156 Yam, Roving, and Cloth, High Modulus, Organic Fiber MIL-Y-83370A (AF) Yams, Graphite, High Modulus, Continuous Filament MIL-Y-83371 (AF) Graphite Fiber Resin Impregnated Tape and Sheet, For Hand Layup MIL-G-83410 (AF) Rods, Pultruded, Graphite Fiber Reinforced, Processing of MIL-R-8712OA (AF) Fabric, Graphite Fiber MIL-F-87121A (AF) Graphite, 1000/3000 Filaments MIL-Y-87125A(AF) HPLC of PMR-15 Polyimide Resin and Prepregs MIL-STD-368 MIL-STD-373 Transverse Tensile Properties of Unidirectional Fiber/Resin Composite Cylinders Transverse Compressive Properties of Unidirectional Fiber/Resin Composite Cylinders MIL-STD-374 In-Plane Shear Properties of Unidirectional Fiber/Resin Composite Cylinders MIL-STD-375 *MIL-STD-2031(SH) Fire and Toxicity Test Methods and Qualification Procedure for Composite Material Systems Used in Hull, Machinery, and Structural Applications Inside Naval Submarines
MIL-Y-1140H MIL-C-8073D
* Military Standard
Table B.6 NASA Standards for composites NASA RP 1092: Standards Tests for Toughened Resin Composites, July 1983 NASA RP 1092 defines five standard tests (STs) for graphite/epoxy composite laminates: ST-1: Compression after impact ST-2: Edge delamination ST-3: Open-hole tension ST-4: Open-hole compression ST-5: Hinged double cantilever beam NASA RP 1142: NASA/ Aircraft Industry Standard Specification for Graphite Fiber Toughened Thermoset Resin Composite Material, June 1985
Non-US Standards in use 1065 Table B.7 FAA Advisory circulars for composites AC 20-10A Composite Aircraft Structures AC 21-26 Quality Control for the Manufacture of AC 145-6
Composite Structures Repair Stations for Composite and Bonded Aircraft Structure
Table 8.8 National aerospace standards for
composites NAS 990
NAS 999
Composite Filament Tape Laying Machine - Numerically Controlled Non-Destructive Inspection of Advanced Composite Structures
conduct research and development on emerging composite materials including characterization, screening, fabrication, and inspection of materials or structures made from the materials. The primary focus of CMC is on screening testing of emerging composite materials. The CMC effort generates a standardized database of consistent properties for advanced composite materials which can be shared between member companies (Table B.9). CMC subcontracts to third parties the material procurement, test specimen fabrication, inspection, test, data analysis, and documentation of results. All tests and data are conducted in accordance with CMC-approved procedures to produce a consistent and standardized database necessary for comparative assessment of material properties. Physical and mechanical properties testing is performed, including lamina and laminate evaluations under tensile,
Table B.9 CMC Member companies
Dow Chemical General Electric Northrop Grumman Lockheed Martin UTC /Sikorsky Loral Vought Rohr Industries
compressive, and shear loadings at selective hygrothermal test conditions. CMC also works with other national organizations to promote composites standardization. B.4 NON-US STANDARDS IN USE B.4.1 AECMA STANDARDS
The European Association of Aerospace Manufacturers (known as AECMA in Europe) is developing various standards for carbon fiber composites for aerospace applications. Table B.10 lists proposed European Norm (EN) standards for determining the physical and mechanical properties of carbon fibers and their composites. B.4.2 JAPANESE INDUSTRIAL STANDARDS (JIS)
Standardization efforts in Japan are managed by the Standards Department under the Ministry of International Trade and Industry (MITI).The Japanese Standards Association is in charge of publishing Japanese Industrial Standards (JIS).Table B . l l lists Japanese standards for carbon fibers and their composites. B.4.3 GERMAN (DIN) STANDARDS German aerospace specifications for carbon, aramid, and glass fiber polymer composites are listed in Table B.12. The DIN specifications and standards are published by the German Institute for Standardization, which is the ’Deutsches Institut fur Normung’ (DIN) in German. B.4.4 INTERNATIONAL (ISO) STANDARDS
The International Organization for Standardization (ISO) Technical Committee 61 on Plastics, Subcommittee 13 on Composites and Reinforcement Fibers is the body which is developing international standards for polymer composites. The USA is represented at international meetings through the American
1066 Specifications and standards for polymer composites Table B.10 AECMA Standards for carbon fibers and their composites
AECMA prEN2557 AECMA prEN2558 AECMA prEN2559 AECMA prEN2560 AECMA prEN2561 AECMA prEN2562 AECMA prEN2563 AECMA prEN2564
Carbon Fibre Preimpregnates, Test Method for the Determination of Mass Per Unit Area Carbon Fibre Preimpregnates, Test Method for the Determination of the Percentage of Volatile Matter Carbon Fibre Preimpregnates, Test Method for the Determination of the Resin and Fibre Content and the Mass of Fibre Per Unit Area Carbon Fibre Preimpregnates, Test Method for the Determination of the Resin Flow Unidirectional Laminates Carbon-Thermosetting Resin Tensile Test Parallel to the Fibre Direction Unidirectional Laminates Carbon-Thermosetting Resin Flexural Test Unidirectional Laminates Carbon-Thermosetting Resin Test Method, Determination of Apparent Interlaminar Shear Strength Carbon Fibre Laminates, Test Method for the Determination of the Fibre and Resin Fractions and Porosity Content
Table B . l l Japanese standards for carbon fiber composites
Table B.12 German specifications for polymer composites
Testing Methods for Carbon Fibers Testing Methods for Carbon Woven Fabrics Testing Methods for Prepreg, Carbon Fiber and Epoxy Resins Testing Method for Tensile Properties of Carbon Fibre Reinforced Plastics Testing Methods for Flexural Properties of Carbon Fibre Reinforced Plastics
DIN 29965 Aerospace; Carbon Fibres, High Performance Carbon Fibre Filament Yams, Technical Specification DIN 29971 Aerospace; Unidirectional Carbon Fibre-Epoxy Sheet and Tape Prepreg, Technical Specification DIN 65090 Aerospace; Textile Glass, Preimpregnated Filament Glass Cloth for E-Glass (Prepreg),Technical Specification DIN 65426 Aerospace; Aromatic Polyamide Part 1: (Aramid) - Preimpregnated Woven Fabric, High-Modulus Filament Yam (Prepreg);Dimensions, Masses DIN 65426 Aerospace; Aromatic Polyamide (Aramid)- Woven Filament Fabric Part 2 Prepreg from High-Modulus Filament Yam and Epoxy Resin, Technical Specification
JIS R 7601 JIS R 7602 JIS K 7071 JIS K 7073 JIS K 7074
National Standards Institute (ANSI). Table 8.13 lists draft and published (ISO) standards for composites. B.5 PROPRIETARY SPECIFICATIONS
This appendix does not include the many composite material and process specifications which have been developed a n d are used by individual private sector companies (i.e. primes, fabricators and suppliers). Most, if not all, of the companies which either produce or use composites have some proprietary specifications for composite materials and processes. It is known that proprietary specifications con-
stitute a substantial data base which cannot be included herein. Please note that this appendix addresses only polymer matrix composites. Specifications and standards for other types of composite materials (MMC, CMC, and C/C) are still in early stages of development and may be either subject to export control or proprietary.
Proprietary spec$cations
1067
Table 813 International (ISO) standards for polymer composites
CD 1268
Fiber Reinforced Plastics - Test Plates Manufacturing Methods -Part 1- General Conditions
CD 3341
Textile Glass - Yams - Determination of Breaking Force and Breaking Elongation
CD 3374
Textile Glass - Mats - Determination of Mass Per Unit Area
CD 4605
Reinforced Products - Woven Fabrics - Determination of Mass Per Unit Area
CD 14127 Composites - Determination of Resin, Fiber and Void Content for Composites Reinforced with Carbon Fiber CD 15024 Polymer
Determination of Mode I Delamination Resistance of Unidirectional Fiber Reinforced Laminate Using the Double Cantilever Beam Specimin
CD 15034
Plastics - Prepregs - Resin Flow
CD 15040 Plastics - Prepregs - Gel Time CD 15310
Reinforced Plastics - Determination of In-Plane Shear Modulus by Plate Twist Method
DIS 3374
Reinforcement Products - Mats and Fabrics - Determination of Mass Per Unit Area
DIS 5025
Textile Glass - Woven Fabrics - Determination of Width and Length
DIS 14126 Fiber Reinforced Plastic Composites - Determination of Compressive Properties in the InPlane Direction FDIS 1889 Reinforcement Yarns - Determination of Linear Density FDIS 1890 Reinforcement Yarns - Determination of Twist FDIS 3344 Reinforcement Product - Determination of Moisture Content FDIS 12114 Fiber-reinforced Plastics - Thermosetting Moulding Compounds and Prepregs Determination of Cure Characteristics FDIS 11667 Fiber-Reinforced Plastics - Moulding Compounds and Prepregs - Determination of Resin, Reinforced Fiber and Mineral Filler Content - Dissolution Methods FDIS 12115 Fiber-Reinforced Plastics - Thermosetting Moulding Compounds and Prepregs Determination of Flowability, Maturation and Shelf Life IS0 2559
Textile Glass - Mats (made from Chopped or Continuous Strands) Basis for a Specification
IS0 3605
Textile Glass - Rovings - Determination of Compressive Properties of Rod Composites
IS0 8515 Direction
Textile Glass - Reinforced Plastics - Determination of Compressive Properties in the Parallel to the Plane of Lamination
IS0 10119 Carbon Fiber - Determination of Density
IS0 10120 Carbon Fiber - Determination of Linear Density CD - Committee Draft DIS - Draft International Standard JDIS - Final Draft International Standard
1068 Specifications and standards f o r polymer composites REFERENCES
11. Documentation from Mr. Cecil W. Scheider, President of Composite Materials Characterization, Inc., 28 October 1991. 12. ISO/TC61/SC13/WG14 Letter Correspondence from Mr. Junichi Matsui, Toray Industries (Composite Materials Research Laboratories; Japan), 5 August 1992. 13. ISO/TC61/SC13 Report to ASTM D20.61, Gary Williams, July 1997.
1. National Advanced Composites Strategic Plan, National Center for Advanced Technologies, September 1991. 2. Traceski, Frank T., Specifcations and Standards for Plastics and Composites, ASM International (Materials Park, Ohio), August 1990. 3. Test Standards and Engineering Databases for Advanced Composites, Draft Position Paper by Aerospace Industries Association, January 1992. 4. Advanced Composites Standardization, “White ACKNOWLEDGEMENTS Paper,” Prepared by the Committee for Standardization of Advanced Composite The following individuals reviewed and comMaterials, 4 November 1992. mented on this paper and are acknowledged 5. Annual Book of ASTM Standards, Vol. 15.03, for their constructive comments: Mr. Jerome American Society for Testing and Materials Persh (retired), formerly Office of the Director (Philadelphia, PA), 1991. of Defense Research and Engineering; Dr. 6. S A C M A Recommended Methods, Suppliers of Gary L. Hagnauer, US Army Research Advanced Composite Materials Association, Laboratory (ARL), Materials Directorate; Mr. (Arlington, VA), 1997. 7. SAE A M S Index, Aerospace Material Specifications, Gary Hansen, formerly Hercules Advanced Society of Automotive Engineers (Warrendale, Materials & Systems Company; Mr. Jerome R. * PA), July 1996. Jaeb, Chief Engineer, Structures Technology, 8.’ Department of Defense Index Of Specifications and Boeing Defense & Space Group; Mr. Samuel J. Standards (DoDISS), 1January 1997. Dastin (retired), formerly Director, Advanced 9. Military Handbook 17, Polymer Matrix Composites, Draft Volumes 1, 2 & 3, US Materials, Grumman Aircraft Systems. Department of Defense, 17 June 1991. 10. Communication with Mr. Joseph R. Soderquist, Federal Aviation Administration (FAA), 4 September 1997.
INDEX
Page numbers in bold type refer to figures; those in italics refer to tables
A-glass 134 Abrading, surface preparation 669 Abrasion mechanical 668 particulate fillers 252 resistance, PET 224,233 Abrasive water jet, see AWJ Abrasives machining 606 inslurry 607 types of 607 ABS compressive strength 258 density 258 maximum service temperature 258 shear modulus 259 shear strength 259 tensile strength 258 thermal conductivity 259 Accelerators 59-61,60-2 anhydride-cured system 72 Acceptance values 724 Accumulation, damage 797 Acids, Lewis 60-1 Acoustic emission 849 honeycomb assemblies 849 fatigue testing 811 pressure vessel 472,472 Acoustical properties, fiberglass 135 Acousto-ultrasonics 849 Acrylic acid, and polypropylene 251 ACTS, reflectors 1017,1017 Acute toxicity 8234,824 measuring 824 Adaptability, adhesives 274 Additives low profile 382 pultrusion 515-18 wetting agents 515-16 Adducts, thermoplastic polymers 526 Adherends axial stress 644
bending failures 644 bending stresses 644 composite 630,660,662-3 deformable 633,634 deformations 630 dissimilar 635,660 ductility 630 equal thickness 650 failures 627-9,628,657 fully tapered 646, 647 identical 639 layered 646 loading stress 642 modulus 643 moisture absorbancy 630 rigid 632,633, 634 stiffness 636 unbalance 629,648 stretching elongations 633 tapering 627,629,64451,645, 647, 648, 649, 650, 653, 656, 658 thermal expansion 630,661 thickness 627-9,628,628,65&7, 658 transverse shear 630 weaker matrix 630 Adhesion aramids 26 failure 674-5 Adhesives 255,271,2744,517-18 adaptability 274 aerostructural 651,652 blisters 271 bond line control 274 bonding 127,255,727 aluminum alloy 683 pressure 271 brittle 651 cell-edge 276 co-curing 868,870 cohesive failure 674-5 core 271 cure temperature 661 Cycleweld 255
damage to 656 ductile response 6516,652,657 ductility 627-8,629-30,655-6 elastic-plastic response 649 environmental conditions 630 epoxies 275 failure in 628 fillet forming 271 flow, sandwich 288 heavy liquids 275-6 high temperature 658 joints 610-63 bonded doubler 628 bonding techniques 611 double-lap 628 double-strap 628 durability 631-2 inspection methods 627 manufacturing deficiencies 611 poor bonding 611 scarf 628,629 single-strap 628 single-lap 628 stepped-lap 628,629 stressesin 632 surface preparation 611 tapered single-lap 628 tapered-strap 628 unsupported single-lap 628 light liquids 2754 mechanical 518 nylon 670 pastes 2754 peel strengths 274-5 stresses 628 phenolics 255 blended 275 Plycosite 255 putties 275-6 Redux 255 repairs 869 reticulating films 276 scrim-supported 873
1070 Index self-adhesive skins 276 shear stresses 632-7,633,634, 635,636 strain energy 629-30,653 supported films 276 syntactic foams 275-6 test methods 630 thixotropic liquids 275-6 toughness 274-5 unsupported films 276 urea-formaldehyde 255 urethanes 275 water absorption 668 Adjusting, lamina properties 763 Advanced Communications Technology Satellite, see ACTS Advanced composites applications 23 military 23 Advanced fibers, repairs 877 Advantages AWJ 604 correlators 834 hand lay-up 352 injection pultrusion 497 machining 596 laser 605 ultrasonic 608 preform 438 pull winding 497 R T M 434-5 tapering 646 turning 597 twin-screw extruders 539 vertical pultrusion 495 AECMA standards 1065,1066 Aerospace applications, materials selection 1009 design 709 goal 712,713 equipment 1004-21 Industries Association (AIA) 1063 industry composite tools 593 facing material 255 prepregs 885 material specifications 1062,2063 Aerostructural, adhesives 651,652 AFR7008, chemistry of 81 AFR700B/S2 laminates air aging 94,95 compression strength 93 flex strength 93,95 shear strength 93 tensile strength 93 weight loss 94 AFRF 605,1024
Agglomerates 313,314 Aging air 85, 90 aircraft 860 thermoplastics 127 Air aging 85,90 3F/36F polyimides 85 AFR700B/S2 laminates 94,95 AvimidN 85 Skybond 85 release agents 516 Airbus A320, composite applications 1036,1038 Aircraft aging 860 aluminum core removal 862 applications 1022-42 certification requirements 1022-3 components 1029-30,1029 current prices 1024,1026 damaged protective coatings 859-60 design 878 process 10245,1026 requirements 713,1022-3 effect of contaminants 863 environmental exposure 860-1 face sheet materials removal 864 face sheets removal 862 fixed wing, current prices 1024, 1026 fluid contamination 859 flushing contaminants 864 honeycomb core removal 862 impact damage 858-9 interim repairs 859 leak paths 859-60 maintainability 1024 maintenance 878 metal face sheets removal 862 metal repairs 857 minor damage 859 moisture barrier removal 863-4 non-metallic core removal 862 open core evacuation 865 production 840-1,841 costs 1024 reinforced plastics 1023 repairs 8574,857-80,1024 planning 861 skin damage removal 861,861 skin penetration 858,859 speed tape repairs 859 stabilizer design requirements 710 preliminary materials 711 structural applications 1027-8, 1027,1028
thermal cycling 818 through-the-face, sheet evacuation 865 water evacuation 863 water ingestion 859 wet lay-up repairs 865-7 Aircraft Energy Efficiency (ACEE) 1035 Alignment, graphitization 191 Aliphatic amine-cured ("ETA) 65 amines, health effects 832,833 content 75 polyether triamine (APTA), viscosity 54 system 70 Alkaline cleaning, metal surfaces 871 degradation 1000 All purpose vehicles 908,908 Allco 59 Alternative, tooling 449-50,449,450 Altex 309,309 Alumina 249-50 composites 325 grit, light grit-blasting 675,676 silicate boria modified 309,320 . strength 309 Young's modulus 309 trihydrate 245 Aluminosilicate, staple 160 AlUminUm alloy adhesive bonding 683 linear thermal expansion 705 core removal 862 CTE 558,589 density 170,558 elastic modulus 170 fiberglass cloth repairs 877 foil 820 fracture elongation 170 oxide fibers 27 particulate fillers 250 rivets 668 specific modulus 1049 strength 1049 surface preparation 871,876 tensile strength 170 tension-tension ratio 1049 thermal conductivity 558 America's Cup yachts 922 Amine curing 49-50,51 epoxy resins, aliphatic 70 Amines, aliphatic 53,54 Amino resins, health effects 834
Index Aminolysis, degradation 888 Amoco 58 anhydride curing agents 56 Analysis composites 736-57 cost of 736 equilibrium equations 737 Ancadride MHHPA (Ciba-Geigy) 57 Anchorage 987 grounded 986 prestressed concrete 9 8 5 6 Angle of impact, meteoroids 815 Angle ply 741 Anhydride curing 50 agents 56-9,56-9 health effects 832, 833 high temperatures 50 Anhydride-cured system 72 accelerator 72 cure cycle 72 curing agent 72 heat distortion temperature 72 maximum stress 72 modulus of elasticity 72 pot life 72 solvent absorption 72 tensile properties 72 viscosity 72 Anis0tropic bodies, thermoelastic behavior 470-1 composites 797 materials 687, 688 material constants 688 pitches, stabilization 178 Anisotropicity, PAN 185 Anisotropy crystal 334,335 GFRP 1006 PET 224 Annealing, and coating thickness 297 Anodizing 667 aluminum alloys 668 low voltage 871 non-tank 871 Antennas 932-3 Antimony oxide 244 highcost 244 synergism 244 Antistatic agents, fiberglass 146-7 Appliances, reuse 901 Applications 3,270,931-56 aircraft 11 antennas 932-3 aramid fibers 216,217-20,222-3 AS4/PEKK 127-8 battery casings 941
biomedical, thermoplastics 127--8 boron 163, 265,166 bridges 9324,933 business equipment 934,934 cable cars 934-5 carbon fibers 196-7 casting 38 ceramic fibers 163,165, 166 CMCs 935 consumer products 946 cooling towers 936-7 drive shafts 936-7 electrical 937,937 electronics 937,937 firefighter breathing apparatus 937-8 flywheel mechanical battery systems 938,938 fuel tanks 939-40 glulam beams 940-1 grating 936 health 9434,944 high silica 163, 165, 166 high-pressure tubing 941 high-speed train brakes 941 hydrogen fuel storage 941 industrial equipment 951 ladders 936 laptop computers 942 lighting poles 948-9, 948 loom components 942 maglev train guideways 942-3 marine 916-28 medical 9434,944 metal surfaces 871 MMCs 9 4 4 5 mobile storage 945 molding compounds 945-6 musical instruments 946 non-aerospace 935 O-rings 953 oil and gas 946-7,947 PBOfiber 238 PET 230 pick-up trucks 936-7 pipeline rehabilitation 947-8 piping 936 PMR-15 96 ports and harbors 949 power poles 948-9,948 quartz 163,165,166 railways 950 rolling stock 949-50 rolls and air shafts 950-1, 951 safety 9434,944 seals 953 SMEs 952 structural 709 tunnels 950
1071
uranium enrichment centrifuges 953-4 valves 954 wind turbine blades 954 Aramid 23,26,425 adhesion 26 composites 210,210 compression 26 costs 717 density 26 fibers 202-22 applications 216,217-20, 222-3 athletic shoes 222 attributes 358 availability 216,217 by-products 222 chemical properties 213-15, 214 choice of resin 222 composites 207 compressive properties 221 costs 223 creep resistance 216,222 cutting 222 definition 202 design of 216,221-2 electrical insulators 215 electrical properties 215 environmental properties 213-15,214 equilibrium moisture content 215,241 failure 210 fatigue resistance 221,222 flame resistance 205,207 forms 216 handling 222 health effects 835 length-to-diameter ratio 204 limitations 207 machining 222 manufacture 2 0 3 4 mechanical properties 207-13, 208-9,209,210,211,212 melting 205 point 204 microbuckling 207 moisture absorption 221 opaque 221 optical properties 215 physical properties 2057,206 pricing 216 properties 20515 reinforced plastic, see AFW safety 222 sandwich construction 222 self-screening 215 solvents 204
1072 Index specific stiffness 207 strengths 207 strength 205 to weight ratio 223 structure 20P5 temperature resistance 205 tensile modulus 715,716 stiffness 209 strength 209 thermal properties 205-7,206 resistance 205 thermoplastic matrices 222 toughness 221 treatments 215-16 twisted 215-16 UV absorption 215,221 water absorption 215 weave 222 weight 222 mechanical properties 718 moisture absorption 26 paper honeycomb 269-70 applications 270 price 716 strength retention 230 strengths 718 tensile strength 26 Aramid/epoxy, CTE 589 Aramid/fiberglass 359 Aramid/graphite 359 hybrids 354 Arc jet tests, carbon-carbon composites 347 Archimedean screw 538 Architecture, microstructural 795, 796 Areas of transition 731-2,732 Armos availability 220 elongation at break 208 initial tensile modulus 208 sources of information 221 specific gravity 208 tensile strength 208 Aromatic amine-cured MPDA 65 amines, health effects 832,833 polyamide fiber, see aramid fibei‘S polyester fibers 235-7,235,235 system 71 Aromatization, and carbonization 183 Articulation components 962-3 Artificial blood vessels 963 joints 810
ligaments, PET 234 AS-4 carbon fiber 123 in-plane shear modulus 124 tensile modulus 123 tensile strength 123 A S 4 / PEEK(APC) observed life 806,807 predicted life 806, 807 residual strength 806,807 strength 125 AS-4/PEKK, applications 127-8 Asbestos, health hazards 252 Assembly bowing, GFRP 1014,1014 ASTM methods, cured epoxy resin systems 72 numbers 66’66 standards 1061,1061 Atmospheric electricity 860 Atomic oxygen 813-14,814 destructiveness 813 erosion 814 fluence 813 reaction efficiency 814 Atomite 247,247 ATOTech 61 ATR72 CFRP wing box 1038 composite applications 1038, 1937 Autoclave 22,116,461 consolidation 298,577 cure cycle 84 definition 577 repairs 868 tooling 589-91,590 Autoclave/oven, molding 361 Automated Real-time Inspection System (ARIS) 842 Automated Tool Manufacture for Composite Structures 566 Automation, cost savings 16 Automobiles, reuse 900-1 Automotive body applications 909-10,913 SMC 909 industry design requirements 713 recycling 901-2 surface preparation 669 interiors 912, 913 market, GMT 127 suspension 911 Availability aramid fibers 216,217 Armos 220 Dyneema 231 glass fibers 24 Hoechst Celanese 231
Kevlar 217-18 PET 231 Spectra 231 SVM 220 Technora 220 Tekmilon 231 Twaron 219-20 Vectran 236 Aviation, price of 1024 Avimid N, air aging 85 An7 advantages 604 disadvantages 604 drilling 604 linear cutting 603,603 machining 600,602 milling 604 process 602 traverse rates 603 tumingwith 6 0 s without abrasives 600 AXAF orbiting observatory 970 Axial compressive properties, Spectra 226,226 loading 745-7 orientation, PET 224 stress 613 adherends 644 tensile properties, Spectra 226, 226 Axisymmetric deformation circular plates 752-3 cylindrical shells 7 5 3 4 disks 752-3 AZ epoxy N (AZS Corp) 63 Bag molding 352-77,361,362, 366-72,368,369,370,371 thermal cures 367 Bagging 577 Balance 9,lO-11 Balanced laminates 693 Balls, tooling 565 Barcol hardness recycled PET 894 testing 38 Barium ferrite, fillers 527 Barium sulfate 249 fillers 527 Barrier creams 829-30,830 layers, silicon carbide 292 materials 863-4 Bars bending 785 testing 784 Bases, Lewis 60-1 Basic structural units, see BSUs
Index 1073 Basket weave 140,145 Bathtub manufacture, Japan 889 Battery casings 941 Beams 745-7,746,995-7,996 bending stiffness 997 curved 791 deflection 995,996 displacements along 747 modes of failure 997 properties of 283 shear deformation 748-9’995 strength analysis 308-9 stresses in 747 thin-walled 747-9,748 types 283 sandwich 281 Bearing failure 616,729 load 620-1,620 contact region 620-1 edge distances 621 joints 626 strength 6265,625 joints 618-19 tests 625 stress 619,619 at fasteners 620 Beech Starship 472,627,1022,1039, 1040 Bell mouth 500-1 Bend radius 308 Bending 785,791-2 bars 785 deflections 637 fasteners 626 single lap joint 641 failures adherends 644 joints 617,617 fasteners 625,625 longitudinal 747 nonlinear 747 rings 785,792 skin fracture 751 stiffness 745 strength 791 Kulon 305 testing 446 VMN-4 305 stresses adherends 644 formulas 284 three-point 785,791 andtwisting 3 whole rings 785 3,3’,4,4‘-Benzophenonetetracarboxylic dianhydride (BTDA), melting point 59,76 Benzyldimethylamine (BDMA),
melting point 60 Besfight 2054 properties of 2 70 BF,MEA 59 Biaxial fabrics, non-woven 504 plain-weave 399 stretching 546 Bicycles 1050 carbon fiber frame 1050 frame tubes 1051 handlebars 1051 tube and lug designs 1050-1 Binders fiberglass 147 organic 314 solutions, availability 82 Bioabsorbable composites 961 biocompatibility 962 Biological attack 810-11 Biomaterials 957-64 Bisaleimide resins 30 seenlso BMIs Bisphenol A fumarates 36,39 homopolymer property 102 structure precursor 102 supplier 102 tradename 102 Bisphenol A-based, health effects 832 Bisphenol E homopolymer property 102 structure precursor 102 supplier 102 trade name 102 Bisphenol M homopolymer property 102 structure precursor 102 supplier 102 tradename 102 Blast pressure 6754,676, 677 Bleed-out 367 Bleeder 577 permeability 581 ply, molding 584-6,585 Blistering CFRP 918 isophthalic polyesters 918 marine applications 918,918 orthophthalic polyesters 918 permeation barrier 918 Blisters, adhesives 271 Blow molding 529,532,549 accumulator 549 polymer properties 550 Blown core, NDE 854 BMC 379-80,491-2
commercial 380 compression molding 492 definition 379 fatigue properties 386,387 flexural modulus 386,386 flexural strength 386,386 injection molded 380 IZOD impact 386,386 recycling 890-3 specific gravity expansion 386, 386 tensile modulus 386,386 tensile strength 386,386 thermal coefficient 386,386 BMI 30,99-114 CAI performance 105 compressive modulus 211 dielectric properties 1 0 6 7 edge delamination 111 F-22 fighter 113,113 film adhesives 114 galvanic corrosion 108-9 homopolymers of 99 hydrolysis 212 matrices 106 mechanical strength 121 molding compounds 114 monomers 100,103 suppliers 112 thermal properties 106,209 toughening, diepoxides 105 BMI-DAB 100-1 definition 105 BMI-MDA, definition 105 Boat hulls 1045 Boeing 777 composite structure 1035,1036 Bolt hole 621 tensions, torque levels 626 Bolted joints 728,728,729,1015 edge distance 617,617 Bolting and riveting 517 Bond defects 630-1 failures 627-9, 628 layer 656-8,657 line control, adhesives 274 stresses 640 lateral deflections 640 thickness 631 to adherend, thickness 641 van der Waals 224,226 Bonded doubler, adhesive joints 628 joints 728, 728,1015 cracking 659,659 creep failure 658
1074 Index design 728 durability 658-9 environmental extremes 658 viscoelasticity 631 step lap joints 627 Bonding adhesives 127,255 agent 190-1 fiber and matrix 801-2 fusion 127 hydrogen PPTA 204 inspection 668 interfacial 149 andjoints 374 metals 668 pressure 872-4 adhesives 271 skin/core 927-18 techniques, adhesive joints 611 Bondline adhesive voids, NDE 854 thickness control 872-3 Bone cement 962 Boron 23 applications 163, 165,166 continuous fibers 163 fibers 27 reinforced plastics (BFRP) 1023 forms 161-3 manufacture 156-7 properties 161 trifluoride (BF,) 59 trifluoride-monoethylene amine (BF,MEA), melting point 61 Boron-aluminum, MMC 299 Boron-epoxy CTE 589 stabilizer covers 1030-1 Bosses 392,392 mold making 390 Boundary conditions 588,750 laminates 689 Bowing, GFRP 1013 Braiding 18,33,164,402,402, 413-18,413,414,415,417,436, 437 2-D 413 definition 415 2-step 416 3-D 413 geometric parameters 416 angle, and fiber volume fraction 417 cross section 417,417 definition 413 dry tows 418 horizontal 413,413
interlacing patterns 415,415 Maypole machines 413,414 multidirectional 473 prepreg 418 processing technology 413-15 rackets 1050 sporting goods 1045-6 structural geometry 415-16 tape 418 track and column 414,414,416 tubular 413 Break-out 597 Breakage 302,319 Breaking strength fiberglass 13940,141-4 Kevlar 214 Technora 224 Breather 288 Bremsstrahlung radiation 816 Bridges 9334,933 deck 986,986 enclosures 992-3,994 rehabilitation 988 Brittle adhesives 651 composites 797 fiber breakage 291 materials, ultrasonic machining 605 solids 188-9 Broken fibers 803 buildup 803 Bromine, halogens 43 BSUs 169 Buckling 547, 747,750-2, 752 asymmetric 755 column-type 755 columns 997 cylindrical shells 755-7 flange 998 local 997-8 pressure 756 sandwich 282 Buffalo Color 58 Building construction, recycling 901-2 systems 989,991 Buildup, broken fibers 803 Bulk molding compound, see BMC Bumpers 815 Bundle theory 189-90 Buoys and floats 927-8 Bushings fiberglass 138 platinum 133 tooling 562 Business equipment applications 934,934 shielding 936
Butt tensile test 638,638 Butyl glycidyl ether (BGE), viscosity 61 By-products aramid fibers 222 volatile 48,78 Bypass load 616,619,620 definition 620 joints 626 pure 620,621 ratio 620 C-glass 134 C-scan pulse-echo testing 842,842 ultrasonic through-transmission testing 841, 841 Cable cars 934-5 Cables 927 Calcium aluminoborosilicate 24 carbonate 246-7,247 abundance 243 fillers 527 low shrinkage 243 Mohs hardness scale 243 particulate fillers 243,247 stiffness 243 Calcium-sodium metaphosphate 957 Calendering 532 Carbide, coating 296 Carbon black 245 content 306 fiber 254,234,33543,336,337 see also CF aluminum interaction 297 applications 196-7 in carbon, matrix 337 compatibility of 295,296 competitive prices 25 composites 190-1 IS0 standards 1067 typical compression strength 1055 cost of 196,917 CVD 335 discontinuous 337-8 elastic modulus 169 electrical conductivity 169 properties 1845,185,186 filament directionality 335 from PAN 171,335,336 from pitch 171 health effects 835 heat treatment 335 high modulus 335 h g h strength 335
Index 1075 honeycomb 270-1 laminate, tensile strength 1056 MPa 425 nanoporous 183 noncircular 336 PAN-based 169 attributes 358 pitch-based 25,353 attributes 358 production 25 pyrolysis 335 reinforced composites 358-9 plastics, see CFRP reuse 884 rovings 502 selection 90 shape 335,336,337 shear strength 295 tensile modulus 335 strength 169 thermal conductivity 169 decomposition 25,25 properties 184-5,185,186 torsional strength 295 volume fraction 335 Young’s modulus 295 fiber-epoxy, drilling 599 fillers 527 specific modulus 1049 specific strength 1049 tension-tension ratio 1049 typeT300 503 Carbon-aluminum infiltration 303 layer reinforced 304 magnesium evaporation 303-4 MMC, tensile strength 300 plate reinforced 304 producing 303 rolling 303 soldering 303 Carbon-carbon, rocket nozzles 712 Carbon-carbon composites 333-50 advanced 344 applications 341,349-50 arc jet tests 347 coating 343-4 cure cycle 342 density 333 high temperatures 340,344 impact damage 334 lay-up 342 linear thermal expansion 333 manufacturing 341-4 mechanical properties 344-5,345
melting point 333 modulus of elasticity 333 multiformity 333,333 oxidation resistance 333 pistons 349 powders 343 pyrolysis 341,342 reproducible strength levels 333 stiffness 333 strength efficiency 344 tensile strength 333 TEOS 349 thermal conductivity 333,349 expansion 333,349-50 gradients 333 oxidation 345,347-8 thermophysical properties 348-9, 349 Carbon-carbon part 2-D 340 in oxidizing atmosphere 340 Carbon-epoxy 1031 composites 27 laminates 669 prepreg 558 Carbon-fiberglass 359 Carbon-graphite, quasi-isotropic laminates 13, 23 Carbon-matrix composites 333 Carbonic, properties of 270 Carbonization and aromatization 183 chemical changes 181-3 and effluent loss 183 emissions during 181-2 microstructural changes 183-4, 184 stabilized PAN 181 Carbonized organic composites 33941,340 curing temperatures 33940 impregnation 340 pyrolysis 340 starting material 339 Carbonyls, thermal decomposition 2934,294 Carcinogenicity 824 Cards combs 495 Carpet plots 718-20,719 construction 719 laminate selection 720 lay-up sequence 719 shear modulus 720,721 strength values 719 tensile modulus 719 strength 720, 720 Cartilage 958
Carving bits, honeycomb 289 Cast aluminum, tooling 442 boron, aluminum cable 299 ceramic, CTE 589 lamina, composites 299 Casting, applications 38 Catalysts 49,832 homopolymerization 49 metal coordination 104 Catalytic curing 50-1, 51 Caul plates 14,288-9 design 591 tooling design 590 Cavity pressure 443 CDS 798,798 CE 99-114,104 adhesives 114 composition 208 compression 122 dielectric properties 106-7 edge delamination 221 galvanic corrosion 108-9 hot-wet performance 108 hydrolysis 112,112 of cyanurate linkages 107 mechanical strength 211 moisture absorption 107 monomers 105 prepreg reinforced 113 property of casting 108 reinforcement 111 suppliers 112 tensile strength 108 thermal properties 106,209 Young’s modulus 208 Celion 1054 compressive strength 87 flexural strength 87 fracture toughness 87 shear strength 87 interlaminar 91 tensile modulus 87 strength 87 weight loss 91,93 Cell configurations, honeycomb 262 size, honeycomb 263 Cell-edge, adhesives 276 Celluloid acetate compressive strength 258 density 258 maximum service temperature 258 shear modulus 259 shear strength 259 tensile strength 258 thermal conductivity 259
1076 Index Cellulosic, fillers 527 Central cylinder, satellites 375-6 Centred injection, ports 451,451 Centreless grinder 429 sander 429 Ceramics 307-29 applications 328-9,328 brittle 361 continuous unidirectional 317-18 cutting tool inserts 328,328 densities 307 density 312 electrical conductivity 307 fibers 27 applications 163,165,166 continuous 163 forms 161-3 health effects 835 manufacture 156-7 properties 161 fracture toughness 312 hardness 307 matrices, designing 317-23,318, 319,320,321 matrix composites, see CMCs lamina, fiber fracture 799 materials 311-12 melting point 322 temperatures 307,311 modulus of rupture 312 Poisson's ratio 312 powder processing 312-14,313, 314 reinforcing 307-11,308,308,309, 310,311 strengthening 3224,323 thermal conductivity 307 thermal expansion 312 tolerance to flaws 307 unreinforced 328-9 Young's modulus 307,312 Certification requirements aircraft 1022-3 military aircraft 1023 CF/PEEK 115 CF/PPS 115 CF/PSU 115 CFFW 1023-4 blistering 918 cathodic 918 prepregs 1-23 Chain folding, PET 224 Channeling 443 Char yield, NR-150 82 Characteristic damage state, see CDS Characteristic temperatures
PET 221 polyamide 6,6 221 polyamide 12 122 polybutylene-terephthalate 121 polyesters 122 polyolefin 122 polypropylene 121 Charge pattern 384-5 preparation 384-5 Chassis application 910-11, 923 Chemical composition, fiberglass 147-8 contamination 860 degradation 299,888-9 name 221 oxidation, direct wet 191 precipitation 292-3 properties, aramid fibers 213-15, 214 resistance 38 and crystallinity 120 feldspar 243-4 fiberglass 134 Kevlar 214 nepheline 243-4 PBI 237 PET 224,230 Technora 214 thermoplastics 126 stability 214 vapor deposition, see CVD vapor infiltration, see CVI Chemically resistant, gloves 829 Chimneys 983 rehabilitation 988 Chlorendic 39 anhydride (CA), melting point 58 resins 36,36,43 Chlorinated solvents, health effects 836 Choice of resin, aramid fibers 222 Chopped strand, mat 155,917 Chopped-fiber, reinforced composites 355 Chopper gun 353 Chord, winding 462 Chromic acid anodized, aluminum foil 820 Chronic toxicity 824 testing 824 CIC 381 modified 381 Circuit boards, fiberglass 136 Circular plates, axisymmetric deformation 752-3 Circumferential, reinforcement 461 Civil aircraft
applications 1035-9,1036,1037, 1038,1039 material weights 1027 Civil engineering, pultrusion 519-20 Clamping 501-2 levels 625 pressure 625 joints 618 loss of 626 Clay grades of 244 particle sizes 244 particulate fillers 244 Clearances, fasteners 624, 626 Cleavage 50 asymmetric 50 thermolytic 81 Climb, milling 596 Close packing 403,403 Closure, of mold 442 CMCs 935 ultrasonic machining 605 CNC, lathes 597 CO,, gas laser 605 Co-poly-p-phenylene/3,4'oxydiphenylene terephthalamide 203,203,205 Coarse fillers silica 516 talc 516 Coating aluminum foil 820 by precipitation 293,294 carbide 296 carbon fibers 2924,294,295 chemical precipitation 292-3 electroless deposition 292-3 in gas phase 294-8 in-mold 385 medical applications 957 metal, heating temperature 294 nickel 292 nickel vapor-deposited 819-20 nitride 296 polyurethane 813 protective 819-20,819 silicon carbide 294 thickness 293,294,297 andannealing 297 and deposition time 294 Kulon 306 VMN-4 306 two-layer 300 Cockroft-Walton accelerator 848,848 COD 124 Coefficient of thermal expansion see also CTE aluminum 558 carbon-epoxy prepreg 558
Index 1077 glass-epoxy 558 high carbon cast steel 558 Invar 558 laminates 322 mahogany 558 monolithic graphite 558 steel 558 whiskers 308 Cohesive fracture 674,675 Collision avoidance 484,485 Color, stability 43 Columns 997-9,999 buckling 997,999 crushing 997 failure mechanisms 997 properties 998,999 slenderness 997 testing 999 Commercial resins 101,102,103 stabilization 178 Comminuted polvmers, fillers 527 Compacted plies, Springer’s model 588 Compaction 470 pressure 457 Compliance 195 Compliant ring method 788-9 Components of composites 879 design 709 sizing 712 Composites 166 adherends 630,660 A1-B-C 304 analysis 736-57 anisotropic 797 aramid fibers 207 reinforced 716, 717 beam cost 723 design example 722-3,722, 723 design values 723 bolting and riveting 516-17 brittle 797 cast lamina 299 components of 879 coupling agents 250 damage assessment 880 sources 838 definition 378 density thermoplastics 534 environmental effects 879-80 fabrication 879 failure 193 analysis 880
fiber dominated 795 flywheels 474-5 glass-reinforced 716, 717 graphite-remforced 716,717 health and safety 880 history 3 5 M hybrids 353 inhomogeneous 797 inspection 880 layer, permeability 581 Materials Characterization, Inc (CMC) 1065,1965 panels 909 preventive maintenance 880 processes health hazards 831 safety hazards 831 processing 879 preforms 397 times 353 protective coatings 879 pultruded, outdoor use 504 quality controls 880 reasons for using 3 repair methods 880 repairs, aircraft 857-8 sealants 879 sealing 302 standard tests 880 talc 246 testing 194-6 thermal properties 661 Ti-A1-C 303 tools 56675 aerospace industry 593 ease of preparation 592 low density 592 prepreg 566,567-71 thermal expansion 592 thermal stability 592 wet lay-up 566 toxicological properties 831-7 tubes curing 428-9 fishing rods 427 unidirectional 191 whisker-alumina 320-1 Composition CE 108 continuous ceramic fibers 309 fiberglass 138-9 fibers 258 whiskers 158 Compound structures 456 Compression aerospace applications 26 after impact (CAI) 105 aramids 26 CE 111
hot-wet 108 measuringin 195 molding 22,43,374,384,529, 532,545 composite parts 549 fiber orientation 548 GMT 117 one-dimensional flow 583-4, 584 polyimides 545 PTFE 545 and pultmsion 490 rackets 1049 thermoplastics 116,544-9 two-dimensional flow 584 UHMWPE 545 molds 394,395,395 properties 360 strength AFR700B/S2 laminates 93 marine laminates 920 normalization 721 strength after impact 125, see also impact energy and tension 786-9 TEOS 346 testing 781 buckling 787 flat specimens 787 rings 782,788-9 transverse cracking 787 unidirectional composites 787 Compressive fiber stress, and fiber volume fraction 581 modulus BMI 212 PMR-15 laminates 88 properties aramid fibers 221 testing 38 strength ABS 258 Celion 87 cellulois acetate 258 epoxies 258 fiberglass epoxies 153 honeycomb 2665,266-7 Kulon 305 laminate 616 Nomex 272,273 PBO fiber 238 phenolics 258 plied-yam 151 PMR-15 82,88 polyurethane 258 polyvinyl chloride 258 recycled phenolic 894 S-glass epoxies 152
1078 Index single-yarn 151 skinned molded foam 258 thermoplastics 124 thermosets 124 VMN-4 305 yam distribution 154 stress, surface 321 Compton scattering 845 Computed tomography 846-7,848 advantages 846-7 density map 846-7 Computer codes, joints 624,630 Computer numerical controlled, see CNC Concrete dilation 988 forms 925 linear thermal expansion 705 polymer 984,984 post-tensioning 986,987 prestressed 985-6 rehabilitation 988 reinforcing 999-1000 Condensed core, NDE 854 Conductivity, and modulus of elasticity 186 Consolidation 314,576-94 autoclave 298,577 definition 576 equation 582 one-dimensional 584 equipment selection 577 fiber deformation 576 history 578-86 mechanisms involved in 576 models 57846,581-6 pressure 577 resinflow 576 techniques 576 temperature 577 thick composites 585 thin composites 585 time 585 window 588 tooling materials 577 Constituent materials 22 properties 766-7 Constitutive equations 739,741 stiffness coefficients 7 4 3 4 Construction applications 982-1001 materials selection 982 carpet plots 719 methods 1014-15 mold 447 pultrusion 519-20 Consumer products 946 Contact
areas 184 inverse method 611 region, bearing load 620-1 Continuity condition, resin flow 582-3 Continuous casting 298-9 ceramic fibers 163,309 dry jet wet spinning 204 fiber-reinforcement 22 fibers 156,338,338 alumina-based 156-7 density 162 description 162 diameter 162 elastic modulus 162 filament directionality 338 winding 338 layer interlocking 338 manufacturers 162 multilayer locking 338 reinforced composites 355 specific strength 162 thermal expansion coefficient 162 trade names 162 filament, fiber architecture 401-2 glass, rovings 492 impregnated compound, see CIC reinforcement 502 strand mat 436,503 whisker 156 Contour, variation 372-3 Contoured tape, lay-up 16 Control surfaces, marine applications 924 Coolants, drilling 600 Cooling towers 936-7,993-4,995 Coordinating partial plies, tooling design 590 Copolymerization 34 Core corrosion adhesives 271 NDE 854 edge treatment 277 fatigue, NDE 854 materials 256-7 movement, adhesives 271 node disbonds, NDE 854 plug repair 866 selection, sandwich 284 shaping 289,289 shear stress, formulas 284 sue, sandwich 288 splitting, adhesives 271 Cores density 276
splices 276,278,278 Comers flexure, TEOS 346 injection, ports 451,451 preform 438 Corporate Average Fuel Economy (CAFE) 907 Correction method, tooling 590-1 Corrective Optics Space Telescope Axial Replacement, see COSTAR Correlation, ultrasonic 8434,844 Correlators advantages 834 block diagram 844 PTFE 834 Corrosion barrier 47 galvanic 108-9 resistance 38,42,303 construction 982,983 high-temperature 322 reinforced composites 387 Corrugated configuration, reinforced composites 389 Corrugation process, honeycomb 257 Cortical bones 958 COSTAR 1018-19,1019 costs aramid 717 carbon fiber 917 E-glass 717 glass fibers 24 graphite 717 meta-aramids 205 metals 717 particulate fillers 242 PBOfiber 238 S2-glass 917 count fiberglass 1 4 1 4 plied-yam 151 single-yam 251 Countersinking, and drilling 598-9 Countersunk fasteners 624,624,625,625 head 729 Coupling agents 671 chemical functionability 147 commercial 148 composites 250 effect on mechanical properties 149 fiberglass 146-7 polypropylene and acrylic acid 251 resin interaction 149
Index 1079 silanes 147,250-1 titanates 251 viscosity 250 zirconates 251 with diepoxides 105 eliminating 193-4 and stiffness 702-3 Coupon tests 195,687,688 Coupons configuration 721 standard 721 Crack bridging 320,3234,323 deflection 320,323-4,323 formation 469 growth, fatigue 811 initiation 124,632 opening displacement, see COD propagation 124 Cracking bonded joints 659,659 matrices 319 Crashworthiness 3 Cratering 814 Crazing 343 Creel bookshelf 492 horizontal feed 494 loaded multiple spindle 494 mat/fabric 494 package positions 493 pultrusion 492-5 Creep effects 126 environmental 1000 failure bonded joints 658 hot-wet conditions 659 fatigue 631 homopolymers 212 low temperatures 800 para-aramids 212 resistance aramid fibers 216,222 PET 227 Spectra 227,228-9,230,232 rupture 797,800,806 time-dependent 800,800 viscoelastic 797 Cresyl glycidyl ether (CGE), viscosity 62 Crimp 416 Critical bend radius, continuous ceramic fibers 309 energy, failure criteria 795 load, orthotropic plates 751 processing temperature 326 shear resistance, failure criteria
795 Crooked chain, meta-aramids 205 Cross section constant 489 shape of 461 Cross sections, strength 613 Crosslinking 31,48,100 matrix 460 reversible 887-8 rotational molding 552 Crowfoot satin, weave 145,145 Crushed core NDE 854 sandwich 282 Crystal, anisotropy 334 Crystalline polymers 118,120 silica 250 Crystallinity advantages of 126 and chemical resistance 120 PET 224 Crystallite melting point, Spectra 26 Crystallization kinetics injection molding 540 rotational molding 552 CTE 698-9,704 aluminum 589 aramid-epoxy 589 boron-epoxy 589 cast ceramic 589 compatibility 557 electroformed nickel 558 fiberglass-epoxy 589 graphite-epoxy 589 high-temperature cast epoxy 589 iron (electroformed) 589 M401/F854 704 nickel (electroformed) 589 shrink factors 557-8 silicone rubber 589 tool steel 589 tooling 556-7,557-8 urethane board stock 558 Cupping 541,541 polyoxymethylene 541 Cure catalysts 104 control 5869,586,587,588 objective 588 cycle anhydride-cured 72 aromatic system 71 autoclave 83-4 stepped 355 carbon-carbon composites 342 elements 67 no-bleed 673
and resin viscosity 354,355 RTM 433 selecting 585,586 times, tools 559 degree of 587,587 monitoring 67-9,67,68,69 in situ 69 infrared spectroscopy 68 off-line 67-9 part slippage 432 rate 509-10,509,510 epoxy resins 509 polyesters 509 shrinkage 48 temperature 105 adhesives 661 Cured epoxy resin systems, ASTM methods 72 Curing 13,16,37,385 agents 49,513,813,832 amine 5 3 4 , 5 6 6 anhydride 56-9,56-9 anhydride-cured system 72 epoxy resins 509 amine 49-50,51 anhydride 50 autoclave 33 composite tubes 428-9 cyclotrimerization 99-101 dicyanates 100 low temperature 567 non-autoclave 589 ovens 22,428-9 press 22 reaction 49 RTM 4456,445 shrinkage 41 temperatures, carbonized organic composites 33940 thermal 499 time 385 Curved beams 791 pultrusion 489-90 cut-off high pressure water jet 502 saw 502,502 station 502 wastage 502 Cut-outs, in cylinders, GFW 1013 cutting aramid fibers 222 Kevlar 600 plies, surface preparation 870 speed 5967,598,605 tools 329,329 CVD 27,157,294,296,311,338-9, 339 carbon fibers 335
1080 lndex drawbacks 317 isothermal 338-9,339 silicon carbide 157 temperatures 316 thermal gradient technique 339, 339 CVI 316,316,401 forced flow-thermal gradient processing 317,317 isothermal processing 317 Cyanate ester, see CE resins 29-30 Cyanates 358 Cyanurate trimer 99,100 CYCAP 93 chemistry of 81 Cycle times, RTM 433 Cycles to failure, Kevlar 213 Cycleweld 255 Cyclic loading, microcracking 801 Cycloaliphatic amines, health effects 832,832,833 Cylinders filament winding 467 Cylindrical shells 466,752-3,753 axisymmetric deformation 7 5 3 4 buckling 755-7 pressure 756 nonsymmetric deformation 754-5 shear deformation 756 stiffnesses 756 Czochralski method 160 D-glass 134 Dagger drill 599400,599 Damage accumulation 797 assessment composites 880 wet lay-up repairs 866 drivers 801-3 inspection, NDE 854 low velocity impact 839-40 modes 797-800,798,799,800 protective coatings, aircraft 859-60 resistance 801-3 tolerance 107, 222, 794, 794-808 definition 794,794 Darcy’s law 447,581 resin flow 578 DCB, tests 124 DCPD 37,37,39 low cost 37 maleate half ester 37 Debonding, moisture 811 Debris impacts 814-15,814,815
Debulking 567,569,592 periodic 574 Deburring 565 Decking 925 Decomposition temperature Kevlar 206 Nomex 206 PBOfiber 238 Technora 206 Teijinconex 206 Twaron 206 Decontamination, PMC 863 Deep delamination, NDE 854 Deep submergence devices 474,475 Defective fibers 303 Defects, inspection methods 733-4, 734 Defibrillation 215 Definitions aramid fibers 202 autoclave 577 BMC 379 BMI-DAB 105 BMI-MDA 105 braiding 413 bypass load 620 composites 378 consolidation 576 damage tolerance 794,794 denier 241 detailed design 710 durability 794,794 electroforming 591 epoxide 48 equilibrium moisture content 241 extrusion 534 fiber 242 placement 476 hazard 823 heat capacity 532-3 hygrothermal 694 knitting 408 lamina 687 material properties 709 matrix 378 particulate fillers 242 preliminary design 710 prepreg 425 pultrusion 488 pyrolysis 888 reinforcements 378 rheology 527 risk 823 specific heat 532-3 specific modulus 715 steering 482 strength retention 241 tenacity 241 Tex 136,241
textile preforming 397 thermosetting reaction 49 tows 476 toxic 822 viscosity 527 weaving 404 Deflection effect 640 equations 750 fasteners 624,626 formulas 284-5 lateral 644 sandwich 284 Deformable, adherends 633,634 Deformation 398 adherends 630 degree of 448 laminates 691 point 302 Degradation 215,796 aminolysis 889 chemical 88&9 factor 697 fiber length 536 glycolysis 889 hydrolysis 889 methanolysis 889 NDE 838 PET 889 products obtained 889 strength 796 Degreasing, metal surfaces 871 DEH 20 (Dow) 54 DEH 24 (Dow) 54 DEH 26 (Dow) 54 de Havilland Mosquito 686, 686 Delamination 548,5974,598,599, 615,798-9,799,801 atedge 798 helical winding 789 Delta Clipper experimental launch vehicle 974-5,974 Demolding 443,4467 precautions 446 DEN 431 (Dow) 52 DEN 438 (Dow) 52 DEN 439 (Dow) 52 Denier, definition 241 Denitrogenation 183 Densification 314 shuttle parts 343 Density ABS 258 aluminum 270,558 aromatic system 71 carbon (Type T300) 503 carbon-carbon composites 333 carbon/epoxy prepreg 558 cellulois acetate 258
lndex ceramic composites 322 ceramics 307 continuous ceramic fibers 309 continuous fiber 162 E-glass 503 electroformed nickel 558 epoxies 258,511 glass/epoxy 558 high carbon cast steel 558 Invar 558 Kevlar 206,503 Kulon 305,306 mahogany 558 meteoroids 815 and modulus 23 monolithic graphite 558 Nomex 206 NR-150 82 PAN-based fibers 170 phenolics 258 pitch-based fibers 170 PMR-15 82 polyester 511 polyurethane 258 polyvinyl chloride 258 recycled NBC 897 recycled phenolic 894 regrindRIM 895 resin, aliphatic system 70 S-glass 503 SiC/Al,O, 315 skinned molded foam 258 Spectra 503 staple 162 steel 170,558 and strength 23 Technora 206 Teijinconex 206 titanium 170 Twaron 206 urethane board stock 558 vinylester 511 VMN-4 305,306 whiskers 162,308 Dental applications, hydroxyapatite (HA) 958 Dentin 958 Deoxidizing, metal surfaces 871 Deposition electrolytic method 293 time, and coating thickness 294 DER 330 (Dow) 52 DER 331 (Dow) 52 DER 332 (Dow) 51 DER 661 (Dow) 51 DER 732 (DOW) 63 DER 736 (DOW) 63 Derivatives, phenolic 382 Description
continuous fiber 162 staple 162 whiskers 162 Design 32 aerospace 709 allowables 709, 758-77 nomenclature 759,759 processing variables 762-3 testing 758 checklist 712,714 components 709 of composites 18-20,19 considerations 994-1001 data, using 724-7 detailed 710 laminates 686708,697-706 low strain 838 methodology knitting 410-13 nonwoven textiles 419 preliminary 710 process 710 requirements aircraft 713,1022-3 stabilizer 710 automotive industry 723 helicopters 713 industrial pressure vessels 713 marine submersibles 713 recreational 713 rocket motors 713 sailboats 713 satellites 713 team 709,712,714,715 materials supplier 715 values 720-2 verification 733-4733 Designations, continuous ceramic fibers 309 Despooling 479 Detailed design 710, 723-34 definition 710 Determinable values 779 Diacids, and glycols 34-5 Diameter continuous fibers 162 ceramic 309 low 160 staple 262 whiskers 262,308 3,Y-Diaminodiphenylsulfone, melting point 55 4,4'-Diaminodiphenylsdfone (DDS), meltingpoint 55 Diaphragm forming 117 thermoplastics 116 Dibromoneopentyl glycol 43 Dicyanates, curing 100
1081
Dicyandiamide (DICY), melting point 56 Dicyclopentadiene, see DCPD Dicyclopentadienyl bisphenol homopolymer property 102 structure precursor 102 supplier 102 trade name 102 Dielectric constant 95 E-glass fibers 231 PET 224 Spectra 231 thermoplastics 107 thermosetting 107 properties 106-7 strength, Skybond 95 Dielectrometry 68,69 bulk 68,69 Diels-Alder 101 reverse 80 Dies 497 chrome plated 501 forming 497 heating and curing 498-501 inspecting 501 internal profile 499-500 manufacturing 500 multi-cavity 491,500 positioning 498 purging 514 steel 500 temperature control 511-13,512,513 profile 511, 513 stability 514 Diethyl ester diacid derivative 80-1 Diethylaminepropylamine (DEAPA), viscosity 54 Diethylenetriamine (DETA), viscosity 54 Different thicknesses, sandwich structures 744 Diffusion welding, solid stage production 291 Diglycidyl ether of 1,4-butanediol (BDE), viscosity 62 Diglycidyl ether of bisphenol A (DGEBA) crystallization 51 viscosity 51 Diglycidyl ether of bisphenol F (DGEBF) 52 crystallization 52 Diglycidyl ether of neopentyl glycol, viscosity 63 Diglycidyl ether of polypropylene glycol 63 Diluents 61-3,62-3
1082 Index reactive 66,66 Dimensional stability, reinforced composites 387 tolerances 560-1 drawings 731 Dimensions, joints 612,649 DIMOX 315-16,315,316 Dimpling, sandwich 282,284 Direct wet, chemical oxidation 191 Direction dependence 698 Directional metal oxidation, see DIMOX Directionally reinforced molding compound, see XMC Disadvantages AWJ 604 hand lay-up 352 injection pultrusion 497 laser machining 605 preform 438 RTM 434-5 turning 597 twin-screw extruders 539 ultrasonic machining 608 Disbonds inspection 839 NDE 854 surfaces 677 Discontinuous fiber 156 representative volume 804 whisker 156 Discrete, fiber architecture 401,401 Disks, axisymmetric deformation 752-3 Displacements, measuring 779 Disposal see also recycling, reuse nonrecyclables 899-900 and reuse 883-904 Dissimilar adherends 635 Dissipation factor 95 fiberglass 135 thermoplastics 107 thermosetting 107 Distortion temperature, testing 38 Diving equipment 926-7,926,927 Dixie Chemical 57 DMP 30 (Rohm & Haas) 60 DMTA 120,120 Dodecenyl succinic anhydride (DDSA), melting point 57 Domes, contours 467 Domier 328 composite applications 1037, 1039 rear fuselage and fin 1039 Double
cantilever beam see also DCB specimens 124 head, pull winding 496-7 lap adhesive joints 628 balanced 655 joints stress distribution 654 thermal stress 662 shear, joints 624,625 strap adhesive joints 628 joints, tapering 646,647, 648 Doublers 279-80,280 Draft 389,390 zero 389 Drapability, preform 438 Draping simulation 448,448 Drawings 730-3 construction 730-3 dimensions 731 tolerances 731 material description 731,731 thickness 731 Drill cutting parameters 600 dagger 599400,599 fixtures, fabrication 573 fully fluted 600 geometry 599-600 performance 599 templates 5 7 3 4 twist 600 Drilling 597400,598,599 carbon fiber-epoxy 599 coolants 600 and countersinking 598-9 damage 729 glass fiber-epoxy 599 machining 602 secondary 598 Drive shafts 911,93&7 Dry spots, eliminating 453,453 Dry tows, braiding 418 Drying, of composite 811 Dual-shell reflector, Kevlar 1004, 1004 Ductile response 655 adhesives 65145,652,657 Ductility, adhesives 62743,629-30, 655-6 DuPont FP fiber 309 elastic modulus 309 Durability 794-808 bonded joints 658-9 definition 794,794 PET 233 predicting 796
Duramite 247,247 Dust particles, milling 597 Dy 023 (Ciba-Geigy) 62 Dy 027 (Ciba-Geigy) 61 DY 062 (Ciba-Geigy) 60 DY 064 (Ciba-Geigy) 60 Dyeing, PET 233 Dyes 516 Dynamic mechanical analysis (DMA) 816 Dynamic mechanical thermal analysis, see DMTA Dyneema 223 availability 231 elongation at break 225 fibertype 225 pricing 232 specific gravity 225 tensile modulus 225 tensile strength 225 E-glass 24,134,425 composite tanks 985 compression properties 360 constituent properties 766 costs 717 density 503 dielectric constant 231 elastic properties 764-7 elongation at break 503 fiber modulus 360 Kevlar, fiber modulus 360 polyester rods 985 references 774-7 reinforced concrete 984 reinforcement 383 specific gravity 383 modulus 1049 strength 1049 strength properties 764-7 tensile failure 383 modulus 360,383,503,716 properties 360 strength 24,383,503 tension-tension ratio 1049 thermal expansion 383 typical properties 1055 uniaxial strength 194 Young’s modulus 24 Ears 371 Earthquakes 197 Ease of preparation, composite tools 592 ECN 1273 (Ciba-Geigy) 52 ECN 1280 (Ciba-Geigy) 52 ECN 1299 (Ciba-Geigy) 52 Economics, land transportation
Index 9056,906 Eddy current testing 849-50 Edge definition, preform 438 delamination 111 distances bearing load 621 bolted joints 617,617 effects 780,789 turning, reinforced composites 390 Edge-defined film-fed growth, see EFG Effect of contaminants, aircraft 863 Effluent loss, and carbonization 183 EFG 160 Egg crate structure 569,570,572 Eight-hamess satin, weave 145,145, 436,436 Einstein coefficient 531-2 fillers 533 Ejection, part 385 Elastic constants graphite/epoxies 4 Kevlar 210 preform 449 displacement 195 modulus aluminum 170 carbon fibers 169 continuous fiber 162 duPont FP fiber 309 Kevlar 205 Kulon 305 PAN-based fibers 170 pitch-based fibers 170 staple 162 steel 170 titanium 170 W - 4 305 whiskers 162 properties E-glass 764-7 graphite epoxies 768-71 Kevlar 764-7 laminate 764-7 S2-glass 764-7 Spectra 764-7 response 319,655 stress-strain, microcracking 798 symmetry, axis of 779 zone length 658 Elastomer, tooling 14 Elbows 464 Electric discharge machining (EDM) 605 Electrical applications 937,937
conductivity carbon black 245 carbon fibers 169 ceramics 307 equipment, pultrusion 519 insulators, aramid fibers 215 properties aramid fibers 215 carbon fibers 184-5,185,186 fiberglass 135 resistance, reinforced composites 387 resistivity, and microtexture 189 Electrically conductive, particulate fillers 250 Electroformed nickel CTE 558 density 558 mandrel 591 thermal conductivity 558 tooling 591-2,592 Electroforming 591-2 definition 591 Electroless deposition, coating 292-3 Electromagnetic interference, see EM1 Electron diffraction 186 Electronics, applications 937,937 Electroplating 820 Element tapering 624 Eliminating dry spots 453,453 microcracking 1012 Elongation 26 epoxy resins 29 NR-150 82 nylon 6/6 PCI-glass 899 PMR-15 82 PP and granulated SMC 897 recycled NBC 897 recycled PP 895,896 regrindRIM 895 Skybond 86 Elongation at break Armos 208 carbon (Type T300) 503 Dyneema 225 E-glass 503 epoxy 511 Hoechst Celanese 225 Kevlar 205,208,503 Kulon 306 Nomex 209 particulate fillers 252-3 PBO 235 polyester 511 Sglass 503 Spectra 225,503 SVM 208 Technora 209
1083
Teijinconex 209 Tekmilon 225 Twaron 208 Vectran 235 vinylester 511 VMN-4 306 EM1 testing 982 buildings 989,991,991 EMI-24 (Air Products) 60 End effects 780 End grain balsa 917 End-capping 81 Energy equation 586-7 fracture surface 320 release rates 632 Engineering constants IM6/epoxy 699,700,701 laminates 693 parameters, preforms 421 properties 22 Environmental aspects, recycling 902-3 conditions, adhesives 630 effects 810-20,1000-1 composites 879-80 creep 1000 history 810 relaxation 1000 exposure 860-1 extremes, bonded joints 658 properties, aramid fibers 213-15, 214 resistance 3,49, 65, 564 Epi-Rez 5014 (Hi-Tek Polymers) 62 Epon 826 (Shell) 51 Epon 1001 (Shell) 51 Epotuf 37-053 (Reichold) 62 Epotuf 37-057 (Reichold) 61 Epoxides 51-3,51-3 curing agent ratio 64 definition 48 molecules 49 Epoxies adhesives 275 compressive strength 258 density 258,505 elongation at break 505 flexural modulus 505 flexural strength 505 heat distortion 505 laminates, tooling 562 maximum service temperature 258 nitrile rubber modified 275 pultrusion 511 reinforced 196,442
1084 Index resins 2&9,29,30,382,504,832 bismaleimides 6 curerate 509 curing agents 509 cyanate ester 6 elongation 29 flexibility 63 formulation 63-4,64 gel times 510,510 glass content 510 modifiers 105 moisture absorption 29 phenolic triazine (PT) 6 polyimides 6 pull loads 510 rigidity 63 selection of 5 viscosity vs. time 510,510 shear modulus 259 shear strength 259 tensile modulus 505 tensile strength 258,505 thermal conductivity 259 Equations of continuity 582 lamina material 760 of motion 738-9,740-1 inertia terms 739 strain-displacment 739 Equilibrium crack spacing 798 moisture content, definition 241 Equipment selection, consolidation 577 wet lay-up repairs 867 ERL 4206 (Union Carbide) 63 Erosion 813 atomic oxygen 814 Ester formation 34 Esterification 35,38 Etching 667 Ether linkages, Technora 210 2-Ethyl-4-methylimidazole (EMI), melting point 60 Europe, recycling 884 European retrievable carrier 971 Evacuation, PMC 863 Evaluation, strength 737 Exothermic reactions 827-8 Exotherms 68 Expansion process, honeycomb 257 Expendable materials 35940 Exposed surface, voids 432 Exposure alkaline 46 assessment 826 limits 824-5, 826 terminology 826 preventing 827
routes 826 to moisture, PET 227 Extended chain PET fibers 202,223-34 manufacture 223-4 exposure, mechanical properties 215 Extruder die extrusion 535,535 shape 535 Extruders single screw 536 twin-screw 536 Extrusion 529 cf. pultrusion 488-9 definition 534 extruder die 535,535 melt pumping 5365,535 melting 534 orientation 538 plasticating 534,535 solid stage production 291 solids conveying 534,535 thermoplastics 526,534-8 Eye protection 830
3F dianhydride 77 3F/36F polyimides, chemistry of 77 FAA advisory circulars 1063, 2065 Fabric formation 402 Fabrication 31,33 composites 879 drill fixtures 573 first article 372 lowcost 33 marine applications 919-20 methods 762 quality 33 techniques 13-18,14,25,16 Fabrics, orientation 504 Face dimpling, formulas 285 sheet materials removal 864 repair 866-7 wrinkling formulas 285 sandwich 282,284 Facing failure, sandwich 282 material 25.56 Failure 193 adherends 627-9,628,657 analysis, composites 880 aramid fibers 210 bond 627-9,628 characteristics joints 659,660
lap joints 643 criteria 193, 795,802-3,802 critical energy 795 critical shear resistance 795 laminates 695-7 envelopes 696 load, joints 615 mechanisms, columns 997 modes 621,622,797400,798, 799,800,804 properties 194-5 stresses, joints 626 Fairings 1024 marine applications 9234, 923, 924 Fasteners bending 625,625,729 deflections 626 clearances 624,626 countersunk 624,624,625,625 deflections 623,624,625,626 design issues 729 diameter 6234,625 increasing 614 effects 612,624-6,624,625 joint around 612,612 loads 622 mechanical 517 multi-rowed 620 multiple arrays 612 parameters 517 protruding head 624,624 selection 729, 729 tension head 625 Fatigue 811 crack growth 811 low temperature 811 para-aramids 213 properties 386,387 resistance aramid fibers 221,222 PET 230 Technora 213 thermoplastics 115 Feedrates 598 Feldspar 243-4 chemical resistance 243-4 particulate fillers 243-4 refractive index 2434 Female tooling, large power yachts 921 Femoral components 960-1 design 960 development 961 FEP 362 mold releases 362 Fiber alignment 494 aluminum based 159
Index 1085 architecture 398 continuous filament 401-2 discrete 401,401 infiltration 401 integrated 401,401 interlooped 402 laminar 401,401 linear 401,401 planar interlaced 402 areal weight 425 breakage, NDE 838 breakout 599 breaks, NDE 854 bundles 403 coating 166,802 composition 158 continuous 156 definition 242 deformation 580-1 consolidation 576 curve 580-1,581 diameter 156 structural hierarchy 778 direction, tensile strength 804 discontinuous 156 distribution 154, 154 preform 438 dominated, composites 795 flow 527-31 cf. particulate flow 530 fracture 799,799 fragments, inhaling 835 length, degradation 536 low risk 835 manufacturers 158 manufacturing processes 258 modulus 360 orientation 527-31 compression molding 548 injection molding 543,543 preforms 421 packing fraction 412 placement 17,47687,477 definition 476 head 476 inspection 4867,487 machines 478,482 materials 478-9 steering 482 surface geometry 484-6 tooling 479-80 properties 400,400 pull-out 320,3234,323 rayon-based 173 reinforced polymers, see FRP reinforcement 22,23,216,435-6 for reinforcing 166 and resin 425 rovings 492,503
separation 615 shape, carbon fiber 335,336,337 strength, degradation 806 tensioning 457 to resin ratio 509 trade names 258,225 type 225 volume 721 fraction 21,404,412,416,578 3-D fabric 416 and braiding angle 417 carbon fiber 335 and compressive fiber, stress 581 Gutowski's model 588 high 585-6 and permeability 399400, 400 processing window 412, 412 waviness 780 Fiberfrax 160 Fiberglass acoustical properties 235 aircraft industry 136 antistatic agents 1 4 6 7 binders 147 breaking strength 239-40,141-4 bushing 138 chemical composition 147-8 chemical resistance 134 circuit boards 136 composition 133-4,234,13&9 continuous 131 strands 147 count 1 4 2 4 coupling agent 146-7 dissipation factor 135 electrical properties 135,136 epoxies 153 fabric count 140 fibers attributes 358 hollow 134 milled 146 filament designations 133 filament diameter 133 filling yam 1 4 1 4 fire resistance 134 fluted core fabrics 146 heat resistance 134 laminates 356 lubricants 14G7 manufacturing 132 mat 137-8 mechanical properties 135 optical properties 235 platforms 989 proof testing 849
rebars 999 reinforced plastics 23 reinforced polyesters, see FRP reinforcing 38 rovings 136,1467,353 specific gravity 135 spire 991 staple fibers 131 strand 131 suppliers 236 surface resistivity 135 tapes 146 Teglass 134 tensile strength 134 thermal properties 135 thickness 1 4 1 4 three-dimensional fabrics 146 volume resistivity 235 warp yam 2424 water resistance 136 weave 1 4 1 4 weight 2414 yam 138 designation 13940 yield 139-40 Fiberglass-epoxy, CTE 589 Fiberite 99 Fibrous dust 835 reinforcement types 795,796 Filament angle of 456 cross-section, PET 225 diameter 139,148-9 Kevlar 206 Kulon 306 Nomex 206 PET 225 Technora 206 Teijinconex 206 Twaron 206 W N - 4 306 directionality, carbon fibers 335 equilibrium, equation 459 equilibrium 459-61 lay-up 457 shape 206 strength, S-glass 24 on surface 459-61 tension 457 winding 17,17,66,116,117, 456-75,762 applications 471-5 continuous fiber 338 cylinders 467 low costs 463 machine 463,464 marine applications 920 pressure vessels 471
1086 Index rackets 1050 spherical shapes 466-7 sporting goods 1046,1047 wet 762-3 Filed inspections, neutron radiography 848 Fill, yams 407,407 Fillers 382, 516 barium ferrite 527 barium sulfate 527 calcium carbonate 527 carbon 527 cellulosic 527 clay 516 coarse, disadvantages 516 comminuted polymers 527 effect on processing 552 Einstein coefficients 533 fine 516 functional 382 glass 527 ground limestone 382 ground petroleum coke 527 inorganic phosphate-based 957 medical applications 957 metallic oxides 527 mica 527 molybdenum disulfide 527 non-functional 382 pultrusion 516 shape of 531 silica products 527 silicon carbide 527 specific heat 533 Fillet forming, adhesives 271 Filling 301,301,302 mold 442 single-yam 151 yam 140 fiberglass 1 4 1 4 Film impregnation 8 Film-stacking 117-18, 718 Filter cloths, PET 233 Finish 215 S-glass epoxies 152 Fire prevention 367 resistance, fiberglass 134 retardance 36,42-5 halogen inclusion 43 retardants, particulate fillers 249 Firefighter breathing apparatus 937-8 Fishing poles 427, 1045, 1051 FIT-technology 117 Flags bias 425,426 longitudinal 425 Flakes 248
impact strength 249 mica 249 silver 249,250 Flame polishing 157 spread 44 Flammability, evaluating 44 Flank wear 599 Flat cylinders, filament winding 467 specimens compression testing 787 testing 781,786-7 tension 786-7 tape, lay-up 16 Flat-layer, microstructures 177 Flaws, planar 840 Flex modulus NR-150 82 PMR-15 laminates 88 Skybond 86 Flex strength AFR700B/S2 laminates 93,95 Celion 87 NR-150 82 PMR-15 laminates 88 Skybond 86 Flexibility 161 Flexible backbone polyester resin 381 mold wall 444 Flexural modulus BMC 386,386 epoxy 511 LPMC 386,386 nylon 6/6 PCI-glass 899 polyester 511 polyetherimide 545, 546 polysulfone 546 PP and granulated SMC 897 recycled NBC 897 recycled PET 894 recycled PP 895,896 recycled SMC 890,892,893 regrindRIM 895 SMC 386,386 vinylester 517 ZMC 386,386 properties Spectra 226,226 testing 38 strength BMC 386,386 epoxy 511 fibreglass epoxies 153 LPMC 386,386 marine laminates 920 nylon 6/6 PCI-glass 899
plied-yam 151 polyester 511 recycled PET 894 recycled phenolic 894 recycled SMC 890,892,893 S-glass epoxies 152 short fiber mat 401 single-yam 151 SMC 386,386 unidirectional tape 401 vinylester 511 woven laminates 401 yam distribution 154 ZMC 386,386 stress PP and granulated SMC 897 recycled PP 895,896 Flexure TEOS 346 testing 195 Float 150 Floc 216 Flocking lay-up, short fibers 337,338 Flow measurements 357 andrigidity 67 Fluid contamination, aircraft 859 Fluids aircraft 812 automotive 812 methylene chloride 812 Newtonian 527,528 Fluorine, toxicity 362 Fluoroethylene propylene, see FEP Flushing contaminants 864 Fluted core fabrics, fiberglass 146 Flywheel mechanical battery systems 938,938 Flywheels 474-5 Foam adhesive voids, NDE 854 core materials 256-7 cores for fiberglass 256 injection 256 insulation 256 in place system 257 polystyrene 256-7 PVC 257 radar transparency 256 shear strength 257 Folding 547 low-cost 117 thermoplastics 216 Forced flow-thermal gradient processing, CVI 317,317 Foreign materials NDE 854 X-ray imaging 844 Formability
Index jamming angles 399 preforming 398-9 weft knitted fabrics 399 woven glass fabrics 399 yarn slippage 399 Forming dies 497 fabrication 497 materials used for 497 Forms aramid fibers 216 boron 161-3 ceramic fibers 161-3 high silica 161-3 quartz 161-3 Vectran 236 Formulas bending stress 284 core shear stress 284 deflection 284-5 face dimpling 285 face wrinkling 285 moment of inertia 285 safety factor 285 Formulation diglycidyl ether of bisphenol A hexahydrophthalic anhydride 64 diglycidyl ether of bisphenol A triethylene diamine 64 diglycidyl ether of bisphenol A triethylene tetramine 64 Fortafil 2054 properties of 170 Fountain flow, injection molding 542 Fractionizing plant 887 Fracture 597 effects, isolating 799 elongation 170 fixation devices 961-2 mechanics, joints 631-2 path 324,324 strength 325 toughness alumina composites 325 Celion 87 ceramic composites 322 high silica 156 moisture 811 NR-150 82 PMR-15 82 Silar 325 Tateho 325 thermoplastics 222,224 Free radical generators 104 FRP 38,41 composites, pultrusion 517 corrosion resistant 44 corrosive attack 45 durability 959
fire retardant 42 flame resistant 44 low bending stiffness 960 orthopedic applications 959 permeability 959 pins 959 properties of 46 pultrusion 47 strength of 45 thermal performance 45 total hip arthroplasty 959 Fuel tanks 93940 Fully fluted drill 600 Functional fillers 382 Fungal growth 810-11 Fusion 157 bonding 127 PEEK 127 Future directions, transportation 915 Galvanic, corrosion 108-9 Gap 487 Gas laser applications 605 CO, 605 Gas-sparging 174-5 Gate 541,542 fiber orientation 543 Gating 450-1 Gel coat 567 spinning 223-4 time 37, 66, 357, 446 aromatic system 72 zone 505 Gelation pressure during 508 resins 499 Gelstar Thermal Analyzer 515 General aviation, applications 103940 equations 73845 Geodesic curvature 459 deviation, angle of 460 line 457 Geometry joints 638-51 single lap joint 641 Germany recycling 902 standards 1065,2066 GFRP 1004-21 aerospace applications 2007 anisotropic behavior 1012-14 anisotropy 1006 applications 1026 assembly bowing 1013,1014, 1014
1087
cut-outs in cylinders 1013 development 1016 diffusivity 1008 economics 1020 hygroscopic nature 2006 impact damage 1006 joints 1012,1012 material cost 2006 microcracking 2006 moisture effects 1007-10 new materials 1019-20 peel strength 2006 predictions 1019 properties 1005 undesirable 1006 springback 1013 temperature extremes 1004-21 warping 1013,1013 woven broadgoods 1024 Gibbsite 245 Glass bottle industry 97 content epoxy resins 510 nylon 6/6 PCI 899 pP/PCI 898 fabrics, knitted 917 fiber-epoxy, drilling 599 fibers 24-5,25 alkaline environment 1000 applications 24 availability 24 cost 24 handling 24 health effects 835 history 131 processing 24 reinforced plastics (GFRP) 1024 reuse 884 S-glass 24 silane coupling 24 toxicity 24 fillers 527 history 131 length, recycled PP 895,896 mat thermoplastics, see GMT reinforced plastic, see GRP rovings 502 to resin, ratio 150-4, 253 transition temperature see also Tg aromatic system 71 moisture 811 RTM 440 Glass/epoxy coefficient of thermal expansion 558 density 558
1088 Index thermal conductivity 558 Gloves 828-9,828 resistant 829 types 829 Glued laminated timber 991 beams 940-1 Glycidyl amines, health effects 832 compounds 831-2,832 ethers, health effects 832 Glycol and diacids 34-5 propylene 35 selection 39 Glycolysis, degradation 888 GMT 115 automotive market 127 compression molding 117 extrusion compounded 118 semi-finished 118,228 Goland-Reissner 641 Golf shafts 429,463,465,698,1051 filament wound 1051 test methods 1051 Grafil 2054 Granulation knife 886 recycling 886-7 Graphite 245,334,334 3-D lattice 334 composites, machining 599 compression properties 360 costs 717 epoxies elastic properties 768-72 references 774-7 strength properties 768-72 fibers 25-6 competitive prices 25 health effects 835 modulus 360 price 716 production 25 reinforced plastics, see GFRP reinforcement 353 specific gravity 383 tensile failure 383 modulus 360,383,715,716 properties 360 strength 383 thermal expansion 383 turbostratic layers 184 Graphite/epoxies 4 CTE 589 elastic constants 4 physical properties 4 strains 4 strength properties 4
thermal data 828 weathering 813 Graphite/polysulfone, weathering 813 Graphite/wood, hybrids 354 Graphitization 169,183 alignment 191 large regions 170 Grating 936 Green form 314 Green strength 446 Grill opening panels, SMC 907 Grinding 600 accuracies 600 cryogenic 886 polymer matrix composites 600 silicon carbide wheels 600 surface speeds 600 Grit blasting, surface preparation 870-1 Ground limestone 382 Ground petroleum coke, fillers 527 Growth factors 558 GRP 839 Grumman F-14 1030 Grumman X-29 aircraft 698,698 Guide pins 442 compression molds 395 Gutowski’s model 578,580,581,582, 583,585 fiber volume fraction 588 numerical schemes 586 GY 281 (Ciba-Geigy) 52 GY 6010 (Ciba-Geigy) 52 Gypsum 564 Half-discs, testing 782 Halogens 44 bromine 43 Hammer handles, pultrusion 491 Hand lay-up 352-77,762 advantages 352 applications, aerospace 375 disadvantages 352 large power yachts 921 marine applications 376,919-20 precautions 363 Handling aramid fibers 222 glass fibers 24 PET 233 Hardeners 49,832 Hardness ceramics 307 regrind RIM 895 tooling materials 577 Harrier VTOL 1031-2 Hazard 823 definition 823
fibrous dust 832 HDPE glass-fiber reinforced 549 graphite filled 549 mica flake reinforced 549 HDT 126, 226 recycled PET 894 Health applications 943-4,944 effects aliphatic amines 832,833 amino resins 834 anhydride curing agents 832, 833 aramid fibers 835 aromatic amines 832,833 bisphenol A-based 832 carbon fibers 835 ceramic fibers 835 chlorinated solvents 836 cycloaliphatics 832, 832,833 glass fibers 835 glycidyl amines 832 glycidyl ethers 832 graphite fibers 835 imides 834 ketones 836 phenolics 834 polyamides 832,833 polyaminoamides 832,833 polyurethanes 834 thermoplastic resins 834 hazards asbestos 252 beryllium oxide 252 composite processes 831 particulate fillers 252 and safety, composites 880 Heat application heating blankets 874 repairs 874-7 risks of 875,877 capacity definition 532-3 thermoplastics 5 3 2 4 cleaned, reinforcements 898-9 deflection temperatures, see HDT distortion 41 epoxy 522 polyester 512 vinylester 522 distortion temperature 38 aliphatic 70 anhydride-cured 72 aromatic 71 lamps, wet lay-up repairs 865, 866 resistance, fiberglass 134
Index 1089 treatment, carbon fibers 335 Heated curing dies 500-1 design 500 Heating blankets repairs 869 wet lay-up repairs 865 radio frequency 499 rate 181,499 single zone 511, 513 Heavy liquids, adhesives 2 7 M Heel blocks, compression molds 394 Helical winding 460 delamination 789 reinforcement 461 Helicopters applications 1040 current prices 1026 design requirements 713 rotor blades 1024 Heloxy WC-63 (Wilmington Chemical) 62 Hercules 3501-6 epoxy resin 579,579 IM7 fiber, thickness 478 HETacid 36 Hexachlorocyclopentadiene, see HET Hexafluorobisphenol A homopolymer property 102 structure precursor 102 supplier 102 trade name 102 Hexafluoroisopropylidene,bridging 77 Hexahydrophthalic anhydride (HHPA), melting point 57 High carbon cast steel CTE 558 density 558 thermal conductivity 558 local stresses 729 modulus, carbon fibers 335 pressure water jet, cut-off 502 silica applications 163,165,166 continuous fibers 163 forms 161-3 fracture toughness 156 manufacture 156-7 properties 161 strength carbon fibers 335 molding compound, see HMC PET 233 reinforced composites 387 temperature adhesives 658 applications 818
carbon-carbon composites 344 cast epoxy, CTE 589 resins 818 sandwich 288 High-pressure tubing 941 High-speed train brakes 941 Higher temperatures, and tensile strength 182 History consolidation 578-86 pultrusion 488 recycling 883-4 HMC 381 Hoechst Celanese availability 231 elongation at break 225 fibertype 225 pricing 232 specific gravity 225 tensile modulus 225 tensile strength 225 Hole fittings, tooling 562 Hollow fiber, fiberglass 134 Holography 851-3,852 laser interferometric 852 phase locked loop 852 Homopolymerization 40,50-1, 51, 511 catalysts 49,59, 60-1 Homopolymerized BF,MEA 65 Homopolymers, creep 212 Honeycomb 257,260-71 aluminum 268 alloy compressive strength 264-5 plate shear modulus 264-5 plate shear strength 264-5 alloys 268 thermal resistance 260 aramid paper 260,269-70 assemblies, acoustic emission 849 carbon fiber 270-1 carving bits 289 cell configurations 262 cell shape 261-3,262 cell size 263 compressive strength 264-5, 266-7 core shear strength 261 corrugation process 257 defects, X-ray imaging 845 density 261,261,262 expansion process 257 glass-reinforced 268 applications 268 compressive strength 266-7,
269 plate shear modulus 266-7, 269 plate shear strength 266-7, 269 Kevlar 271,272,273 paper 271 metal, roll-forming 289 non-metallic, thermal resistance 260 panels, repairs 871-7 paper 263,268 plate shear modulus 264-5, 266-7 plate shear strength 264-5,266-7 specimen geometry 263 stainless steel based 257 test method 263, 263 thickness 263 titanium based 257 Hoop stresses 621 winding, reinforcement 461 wound, rings 789 Horizontal drawing 299 tape wrapper 428,428 Hot air blowers, wet lay-up repairs 865 extrusion compression 298 MMC 291 oil jackets 499 pressing 314 spots, particulate fillers 242 Hot-wet service 99 HRDI optical bench 1017,1018 HS carbon composites shear modulus 1057 tension 1057 HT 972 (Ciba-Geigy) 55 HT 976 (Ciba-Geigy) 55 HT 9720 (Ciba-Geigy) 55 Hubble Space Telescope 967,969 Humidity 3 Humphrey Chemical 57 H Y 906 (Ciba-Giegy) 57 Hybridization, PET 233 Hybrids 795 aramid/graphite 354 composites, reuse 883 graphite/wood 354 materials 359 Spectra/graphite 354 Hydraulic ejection, part removal 446 system, testing 782 test technique 782, 788
1090 Index Hydrocarbons 37 Hydrocodes 815 Hydrogen fuel storage 941 reduction 161 Hydrogenation 179 Hydrolic stability 38 Hydrolysis BMI 112 CE 112 degradation 888 resistance 41 Hydrophobic, PET 230 Hydrous aluminosilicate 244 Hydroxyapatite (HA) 958 dental applications 958 glass-reinforced 958 Hygroscopic nature, GFRP 1006 Hygrothermal definition 694 effects 6 9 4 5 load 695 properties 96 Kevlar 222 Hypersonic vehicles, materials for 973
ICBM equipment 375 Identical adherends 639 IM6/eqoxy englneering constants 699,700, 701
longitudinal CTE 699 extension 702 strain 702 tensile modulus 699 Poisson's ratio 699 shear modulus 699 stiffness coefficients 702, 703 strength ratio 701 transverse CTE 699 extension 701 modulus 699 strain 703 volume fraction 699 IM-7/PEEK, strength 125 Imidazoles 104 Imides, health effects 834 Impact damage aircraft 858-9 carbon-carbon composites 334
GFRP 1006 tap testing 858 energy, thermoplastics 225 resistance, moisture 811
strength aliphatic system 70 flakes 249 testing 38 TEOS 346 Imperfections 780 Implants, biologic response 960 Impregnation carbonized organic composites 340
powder 118 prepolymer 118 Impregnators, marine applications 919-20
In-mold coating (IMC) 385 In-plane shear Kevlar 212 methods, testing 783 modulus, AS-4 carbon fiber 124 Incineration 888 nonrecyclables 899-900 Inclination angle, yams 407 Inclusion of particulates, polymers 528
Inconel 329 Industrial equipment 951 hygiene 825-30,827,828,829 order of priorities 828 pressure vessels, design requirements 713 Industry, pulp and paper 45 Inert,PET 230 Inertia 3 Infiltration carbon/aluminum 303 fiber architecture 401 improving 300 isothermal gradient 338 of preforms 299 pressure gradient 338 pressure pulsation 338 spontaneous 300 technology 300 under pressure 299-300 vacuum 299-300 Inflatable bladders 874 Infrared thermography 850-1, 850, 851, 1059
aerospace 851 marine applications 851 Inhalation, measure by 824 Inhibitors, and initiators 382 Inhomogeneities, NDE 838 Inhomogeneous, composites 797 Initial rise 646, 647 Initial tensile modulus Armos 208 Kevlar 208
Nomex 209 SVM 208 Technora 209 Teijinconex 209 Twaron 208 Initiators, and inhibitors 382 Injection cycle 442 gate 452 molded 380 molding 22,529,532 cavity filling 542 clamp 5 3 9 4 , 5 4 0 crystallization kinetics 540 fiber orientation 543,543 fountain flow 542 particulate fillers 242 plasticating 538, 538 polymer 540 pumping section 538 rackets 1050 resins 122 screws 539 thermoplastics 226,526 ports 442-3 centred 451,451 corner 451,451 pressure 444
RTM 433 pultrusion 497, 498 sensor controlled 452-3 Inplane shear 374 Inspection bonding 668 composites 880 criteria 487, 838 disbonds 839 fiber placement 486-7,487 first article 486 in-service 839 methods adhesive joints 627 defects 7334,734 and nondestructive testing 1059 on-aircraft 849 paint surfaces 839 part 733 post cure 572-3 speeds 842 subsurface damage 839 times 839 ultrasonic 1059 visual 839,839-40 Insulation foam 256 resistance, Skybond 95 Integrated, fiber architecture 401, 401
Inter-yam slip 448
Index Interaction, constituents 801-2 Intercontinental ballistic missile program 353 Interface, fiberlmatrix 123 Interlacing patterns, braiding 415, 415 Interlaminar fracture 682 shear 790-2 TEOS 346 shear strength Celion 91 PMR-15 laminates 88 stresses 738 eliminating 801 tension, TEOS 346 Internal bone fixation 957 friction background 296,296 mold release 383 International Organization for Standardization 1066,1067 Interphase, toughening 802 Invar CTE 558 density 558 linear thermal expansion 705 thermal conductivity 558 Inverse method, contact 611 Ionic spraying 300 Ionizing radiation 816 Iosipescu shear test 790 Iron (electroformed), CTE 589 IS0 standards, carbon fiber composites 1067 Isophthalics 39 polyester blistering 918 marine applications 916 resins 504 resins 35,36 Isothermal processing, CVI 317 thermal gradient, infiltration 338 Isotropic casting, short fibers 337,338 layers, different moduli of elasticity 744 materials 687 metallic tape, winding 468 nonwoven fabric 164 plates 614 Italian, talc 248-9, 248 Izod impact BMC 386,386 nylon 6/6 PCI-glass 899 PP and granulated SMC 897 recycled
NBC 897 PET 894 phenolic 894 PP 895,896 SMC 892,893 SMC 386,386 LPMC 386,386 recycled, SMC 890 ZMC 386,386 Jamming 412,417 angles 399 Japan bathtub manufacture 889 recycling 887,888,903 Japanese Industrial Standards 1065, 1066 Jeffamine T 403 (Texaco)54 Joining 727-30,728,729 selection process 727,728 thermoplastics 127 thermosets 127 Joints adhesive 517-18,61043,62743 bearing load 626 bearing strengths 618-19 bending failures 617,617 bolted 611, 617, 617,1015 bonded 658-9,1015 step lap 627 and bonding 374 bypass load 626 clamping pressure 618 computer codes 624,627,630 design 728 dimensions 612,649 disassembly 611,728 double shear 624 failure 615,618,618,626,659,660 fracture mechanics 631-2 geometry 613-17,613,628,629, 638-51 GFRP 1012,1012 load magnitude 728 materials 618-19,618,619 mechanical response 630 mechanically fastened 61043 advantages 517 cf. adhesive 610-11 metallic 613 multi-fastener 612, 613,619-24, 620,621,622,623 multi-row 616 open hole coupon strength 626 predicting peak stresses 613 pure bearing load 626 scarf 611,623 selecting design 627 single fastener 612, 612,613
1091
single lap 626, 627, 729 single shear 624 bending moments 624,625 step lap 645-6, 645,649, 650, 651, 652,659 strength experimental 616 predicting 615 tests 626 stress analysis 611 structural performance 613 test methods 626 two fastener 623 untapered 624 Kaolin calcined, hardness 244 particulate fillers 244 Kapton 820 Kardos’ model 578,583 Kerf cutting 604,606 width 606,695 Kerimid 99 Ketones, health effects 836 Kevlar 108,472,598,761,1024 availability 217-28 breaking strength 214 charring 815 chemical resistance 214 stability 214 colored 216 constituent properties 766-7 cutting 600 cycles to failure 213 decomposition temperature 206 density 206,503 dual-shell reflector 1004,1004 E-glass 360 elastic constants 210 modulus 205 properties 764-7 elongation at break 205,208,503 filament diameter 206 shape 206 hypothermal properties 212 in-plane shear 212 initial tensile modulus 208 linear thermal expansion coefficient 206 longitudinal compression 212 longitudinal tension 212 machining 600 marine applications 917 melting 815 temperature 206
1092 Index milling 600 modulus and temperature 210, 211 moisture content 206 properties 205 references 774-7 refractive index 215 reinforcement 383 sources of information 221 specific gravity 208,383 heat 206 modulus 1049 strength 1049 strength retention 205,215 stress rupture 212-13,213 tensile failure 383 modulus 383,503 strength 208,383,503 and temperature 210, 211 tension-tension ratio 1049 thermal coefficient 383 conductivity 206 transverse compression 212 tension 212 trimming 600 turning 600 twist 215-16 ultimates 210 weight loss 205 Kevlar/epoxy thermal data 828 weathering 813 Kine1 99 Kinematics, laminates 690-1 Kinking 302 Kneaded molding compound 381 Knitted glass fabrics 917 reinforcements 916-17 Knitting 402,402,408-13,409,410, 411,412 3-D fabrics 409 definition 408 design methodology 410-13 machines 409 multiaxial warp knit, see MWK stitch formation 409 unit cell geometry 411 warp 408,409 weft 408,409 Kortex 108,271 Kozeny constant 579 Kozeny-Carman equation 400,422, 422,579 Kraft paper 268
Kulon 300 bending strength 305 carbon content 306 coating thickness 306 compressive strength 305 density 305,306 elastic modulus 305 elongation at break 306 filament diameter 306 longitudinal CTE 306 specific modulus 305 strand properties 306 tensile modulus 306 tensile strength 305, 306 Labour requirements, pultrusion 489 Ladder polymer 178-9,178 Ladders applications 936 pultruded 518 Lamina 687-9,688 allowables 759-62 definition 687 macromechanics 192-3 material equations 760 properties 760 properties adjusting 763 estimating 761-2 references 772-3, 774-7 three-dimensional 761 two-dimensional 760-1 thickness, structural hierarchy 778 Laminar discontinuities 843 fiber architecture 401,401 Laminated plate theory 689,690-2 structures 585 Laminates advantages of 322 allowables 762 analysis of 11 balanced 194,693 bending 730 boundary conditions 689 carbon-epoxy 669 coefficient of thermal expansion 322 compressive strength 616 cross-ply 322 deformation 691 design 68&708 elastic properties 322, 764-7 engineering constants 693 failure criteria 695-7 failure in tension 696
fiberglass 356 IM6/epoxy 699,700,701 kinematics 690-1 linear bending 749-50 load carrying capability 696 loads on 689,693 macromechanics of 193-4 mechanical properties 717 nonsymmetric 752 off-axis stiffness 689-90 with plies 697 ply angle 689,689 ply stacking sequence 692 quasi-isotropic 9-11,697 reinforced 321-3,321,619 resulting strain state 6 9 3 4 selection, carpet plots 720 skin 694 stacking sequence 616,630 stiffness matrix 693 strength 695,696, 764-7 ratio 696 stress patterns 322 resultants 691-3 symmetric 194,692,749 thermal stresses 322,589 void free 588 Land transportation 905-15 economics 905-6,906 history 9068,907 market growth 905-6,906 Landfill, nonrecyclables 899-900 Lap joints 638-44 abrasive cleaning 872 failure characteristics 643 peel stresses 637-8,637,638 single 640,641 symmetric 639 Laptop computers 942 Large Area Composite Inspection System (LACIS) 842,842 Large diameter, PAN 180,181 Large power yachts blister protection 921 displacement hulls 922 female tooling 921 hand lay-up 921 marine applications 921-2 variable mold 921 Laser beam 601 heated floating zone, see LHFZ machining 605 shearography 853,853 immunity to vibration 853 Lateral cohesion 184 compression tests 449
Index deflections 642 bond stresses 640 Lathes, CNC 597 Lattice fringe imaging 186 structures 743, 744 Lay-up 15,16 carbon-carbon composites 342 contoured tape 16 comer techniques 373 filament 457 flat tape 16 manual 16 molds 5 6 6 7 operation 576 sequence, carpet plots 719 wet 353,355 Layer reinforced, carbon/aluminum 304 Layered, adherends 646 Layers coordinates 740,746 orthotropic 742 LDEF 970-2,970,971,972 composite specimen testbed 970-1 orientation in orbit 971 Lead powders 249 Leaf springs 910,910 Liteflex 910 Leak paths, aircraft 85940 Leakage 456 Legal aspects, recycling 902-3 Length to diameter ratio, aramid fibers 204 whiskers 308 LEO 813-17 Lewis acids 50-1,60-1 bases 50-1,60-1 LHFZ, directional solidification 160 Life expectancy, tools 559 Light aircraft 1024 grit-blasting 668,670,671 alumina grit 675,676 machine 677,677 liquids, adhesives 2754 weight, reinforced composites 387 Lighting poles 948-9,948 Lightweight fillers, particulate 249 Limitations, aramid fibers 207 LIMS 447,453 Lindau 59 Line source 453 Linear bending 745-7
laminates 749-50 cutting, machining 602 elastic response to failure 615 fiber architecture 401,401 thermal expansion aliphatic system 70 aluminum alloy 705 aromatic system 71 concrete 705 Invar 705 steel 705 titanium alloy 705 thermal expansion coefficient carbon-carbon composites 333 Kevlar 206 Nomex 206 Technora 206 Teijinconex 206 Twaron 206 yams, insertion 410,410 Linkages, flexible 75 Liquid injection molding simulation, see LIMS processing 314-16 stage, MMC 291 waste 889 as fuel 889 Lithium alumino silicate (LAS) 318 LMOs 179,182 size of 179-80 Loading conditions 736 history, remaining strength 804 levels, particulates 528 methods 779 selecting 787 sections, test specimen 786 slow cyclic 627 stress, adherends 642 Loads bypass vs. bearing 620,620 carrying capability, laminates 697 direction of 779 distribution 621 enhancement 655 fastener, variable 622 hygrothermal 695 measuring 779 path eccentricity 637,637 single lap joint 637,637 transfer 610 Local molecular order, see LMOs Localised, microbuckling 800 Logging 3 Long Duration Exposure Facility, see LDEF Long-term
1093
loading 213 microcracking 801 use temperature 206 Longitudinal bending 747 compression, Kevlar 222 CTE IM6/epoxy 699 Kulon 306 M401/F854 699 PBOfiber 238 extension, IM6/epoxy 701 ply waviness 432 strain, IM6/epoxy 701 tensile modulus 699 tension Kevlar 212 test 688 Loom components 942 Lot-to-lot variation tensile modulus 725 tensile strength 725 Low bending stiffness, FRP 960 cost parts, RTM 433 density composite tools 592 polyethylene, viscosity 528 K-glass 134 pressure grit blasting 631 molding compound, see LPMC temperature, fatigue 811 toxicity, thermoplastics 115 viscosity, ZMC 380 void, composites 79 voltage anodizing, metal surfaces 871 Low-earth-orbit, see LEO LPMC 381 flexural modulus 386,386 flexural strength 386,386 formulation 384 IZOD 386,386 shelf life 381 specific gravity expansion 386, 386 tensile modulus 386,386 tensile strength 386,386 thermal coefficient 386,386 Lubricants, fiberglass 146-7 M60J tape mechanical properties 718 strengths 718 M401/F854 CTE 704 longitudinal 699
1094 Index transverse 699 longitudinal tensile modulus 699 Poisson's ratio 699 shear modulus 699 transverse modulus 699 volume fraction 699 m-phenylenediamine (MPD) 55,76 McDonnell Douglas Bea AV-8B, material usage 1032 C-17A, composite applications 1033 F/A-lSE/F, material usage 1031 MD-11, composite structure 1036, 1037 Machinery, marine applications 925-8 Machines caterpillar type 502 dual mandrel stations 477,478 fiber placement 478 light grit-blasting 677,677 tools 363-5 Machining 363-5,364,596408 abrasives 606 advantages 596 aramid fibers 222 characteristics 596 closed loop 3645,364 drilling 602 electric discharge 605 graphite composites 599 Kevlar 600 h e a r cutting 602 milling 602 PET 233 requirements 602 turning 602 ultrasonic 605-8 Macromechanics, of laminae 1 9 2 4 Maglev 979 train guideways 943 Magnamite 1054 Magnesium oxide, particulate fillers 250 Magnetic transparency 983 Mahogany CTE! 558 density 558 thermal conductivity 558 Maleic anhydride 35 isomerization 35 melting point 58 Mandrels 429,459,4646,465 basic requirements 465 collapsible 466 cone-shaped 482,483 dissolvable materials 465 electroformed nickel 591 fusible materials 465
graphite 464 instrumented 468 low cost materials 464 metallurgy 429 pressure on 468 puller 428 removable 464-5 removal 466,467 selection 465 spider/plaster 466 surface 591 table rolling 429 tensometric 468 water-soluble sand 465-6 Manufacturers continuous fibers 162 ceramic 309 fibers 158 PAN-based tow 198 pitch-based tow 199 staple 162 whiskers 158,162 Manufacturing 6 aramid fibers 2034 boron 156-7 ceramic fibers 156-7 deficiencies, adhesive joints 611 extended chain PET fibers 223-4 fibreglass 132 high silica 156-7 options 22 quartz 156-7 SMC 381 Marine applications 916-28 adhesives, paste 918 America's Cup yachts 922 blistering 918,918 buoys and floats 927-8 cables 927 concrete forms 925 control surfaces 924 CORS 917-18 decking 925 diving equipment 926-7,926, 927 environmental effects 918-19 fabrication 919-20 fairings 9234, 923, 924 filament winding 920 future developments 928 hand lay-up 919-20 impregnators 919-20 isophthalic polyester 916 Kevlar 917 large power yachts 921-2 machinery 925-8 microbial degradation 919 mine counter measure vessels
920-1 oil platforms 9245,925 phthalic anhydride 916 piping systems 926 pressure hulls 922 propulsion shafting 925 reinforcements 916-17 resins 916 transfer molding 919-20 shipboard armor 925 small boats 921 sonar domes 922 Spectra 917 vacuum bag processing 919-20 vinyl esters 916 bacteria 919 construction 994 fouling 81G11 laminates 920 submersibles, design requirements 713 Market growth, land transportation 9054,906 Mass transit applications 914 Master models 563-6,592 fabricating materials 563 storing 563 Mat 164 chopped strand 138,155 continuous strand 138 fiberglass 137-8 surface 138 Matched die, molding 361 Materials anisotropic 687,688 cost, GFRP 1006 definition 731 density, X-ray backscattering 846 description 731,732 difficult to roll 430-2 fabrication 786 isotropic 687 monolithic 460 preliminary design 715-22 procurement costs 722 properties 736 definition 709 T50 graphite 724, 724 T50/F584 epoxy 726,726 property, equations 760 quality 780 quasi-isotropic 687 selection 45,712 applications aerospace 1009 construction 982 tools 559 specification 722
Index table rolling 429-30 weights 1027 wet lay-up repairs 867 for winding 458-9 Matrix 28 adhesion 215 cracking 319,798 rateof 806 cracks NDE 854 X-ray imaging 844 crosslinking 460 definition 378 formulation 101-5 materials 356-8 ceramic composites 311-12 polyester 28 systems 4 6 moisture absorption 5 thermosetting 28 transfer molding 315 vinyl ester resins 28 Maximum bond peel stress 642,642-3,644 shear stress 642-3,644 service temperature 258 stress anhydride-cured system 72 aromatic system 71 Maypole machines, braiding 413, 414 Measuring displacements 779 loads 779 strains 779 Mechanical fastening 727 joints, design process 729 pressure 874 properties aramid 207-13,208-9,209, 220,211,212,718 carbon-carbon composites 344-5,345 extended exposure 215 fiberglass 235 laminates 717 M60J tape 718 PlOO tape 718 T300 fabric 718 and weave pattern 252 strength 112 tests 778-92 Mechanically fastened joints 610-63 disadvantages 517 repairs 877
resistant, gloves 829 Mechanics 18,18 Median lethal dose 824 Medical applications 9434,944, 95744 coatings 957 fillers 957 polylactids 957 polyorthoesters 957 strength retention 957 Melt extrusion, PET 224 flow index, see MFI impregnation 117,118 infiltration 315 pumping, extrusion 5365,535 spun, Vectran 235 stretching 529-30 Melting aramid fibers 205 extrusion 534 point carbon-carbon composites 333 ceramic composites 322
4,4'-diaminodiphenylsulfone (DDS) 55 PET 225 recycled NBC 897 S-glass 24 temperature see also T,,, ceramics 307,311 Kevlar 206 Nomex 206 Technora 206 Teijinconex 206 Twaron 206 Vectran 235 Mesophase injection 176 pitch, precursors 335,336 Meta-aramids cost of 205 crooked chain 205 manufacture 203-4 Metal matrix composites, see MMC molten 812 repairs, aircraft 857 surfaces 871 Metallic devices, total hip arthroplasty 960 joints 613 oxides, fillers 527 Metallurgy, mandrels 429 Metals, costs 717 Meteoroids 81615
1095
density 815 impacts angle of 815 space 81615,814,815 Whipple-type shield 814-15, 814 size of 815 velocity 815 Methacrylic acid 40 Methanolysis, degradation 888 Methylene chloride 812 4,4'-methylenedianiline (MDA) 55, 76 Methyltetrahydrophthalic anhydride, melting point 59 MFI 529 Mica 2 4 5 4 fillers 527 Microbial degradation, marine applications 919 Microbuckling 799-800,800 aramid fibers 207 localised 800 Microcracking 323-4,323,797-8, 798,801,1011-12,1012 cyclic loading 801 elastic stress-strain 798 eliminating 1012 GFRP 1006 long term loading 801 minimizing 1012 stiffness changes 796 Micromechanics 191-2,796-7 Microspheres 246,248 hollow 246 particulate fillers 246 types of 246 Microstructures 803 flat-layer 177 onion-skin 177 radial 177 random 177 Microtexture and electrical resistivity 189 and tensile strength 188 and Young's modulus 189 Microwaves radome wall 113 testing 853 transparency 109 Migration 577 Military aircraft certification requirements 1023 components 1030-5 applications 914-15 specifications 1062, 1064 Milling 596-7
1096 Index AWJ 604 ball or hammer 886 climb 596 disadvantages 597 dust particles 597 Kevlar 600 machining 602 with polycrystalline diamond 596 recycling 886-7 square pockets 604 Mine counter measure vessels 920-1 Minimizing, microcracking 1012 Minor damage aircraft 859 dent fillers 859 Minor impact resistance, reinforced composites 387,388 Mix viscosity, resin 441 MMC 27,291-306,944-5 alumina/carbon 291 applications 1005 boron/aluminum 299 by hot extrusion 291 carbon reinforced 291 continuous casting 298-9 cross sections 298, 298 liquid stage 291 production 291 recyclability 944 solid stage production 291 solution sedimentation 291 Mobile storage 945 Modulus adherends 643 anddensity 23 of elasticity aliphatic system 70 anhydride-cured system 72 aromatic system 72 carbon-carbon composites 333 and conductivity 186 fiberglass epoxies 253 plied-yam 251 S-glass epoxies 252 single-yarn 252 yarn distribution 254 quasi-isotropic 12 of rupture, ceramics 322 S-glass 24 transverse 12 Mohs ratings 251-2 Moisture 811-12 absorbancy adherends 630 RTM 440 absorption 695,811 aramids 26,221
CE 107 epoxy resins 29 PET 224 PMR-15 82 resins 440 tooling 561 barrier removal, aircraft 8 6 3 4 contamination, X-ray imaging 858 content 206 debonding 811 desorption 817 detectors 858 diffusion, GFRP 1007-8 effects 96 fracture toughness 811 glass transition temperature 811 impact resistance 811 prebond 670 regain, PBO fiber 238 resistance, PBO fiber 238 transverse strength 811 vs. time, P75S/cyanate 1010 Moisturizers 829-30,830 Molded surfaces 392 Mo1ded -in color, reinforced composites 388 threads 392 Molding autoclave 367 autoclave/oven 361 bag 36672,368,369,370,371 bleeder ply 5846,585 blow 529-30,529,532 compounds applications 945-6 BMI 114 compression 127,365, 384,545 enclosed 366 matched die 361 matched metal compression 378-88 negative draft 366 oven/press cure 366 preparations 361-2 pressure 385,545 bag 367 resin transfer 374,492 seealso RTM sheet 374 thermal expansion 365-6,365, 366,593 vacuum 385 bag 361,366,577 waste 374 Moldless construction 257 Molds 441 aluminum 361 backingup 422
cavity design 442 closing 385,442 speed 385 construction 447 design 362 dual steel 361 elements of 541 filling 385,442, 4 4 6 5 pressure during 452,452 simulation 450-1 graphite-epoxy 590 maintenance 361 making 390,390 materials 441-2 metal 441 polymeric composite 441 preparation 442 release 361,362,428,429-30,514 problems and solutions 568 secondary 430 RTMdesign 441 sealing 443 shrinkage nylon 6/6 PCI-glass 899 recycled NBC 897 steel 442 selection 393 stops, compression molds 395 stresses 392-3 temperature control 541 Molecular arrangement 133 characteristics, rotational molding 552 orientation 27 weight, thermoplastics 116 Molybdenum disulfide, fillers 527 Moment of inertia, formulas 285 Monobands, sealed 304 Monofilaments 164,307 reinforcements 311, 322 Monolithic graphite 559,561,562-3 coefficient of thermal expansion 558 density 558 low CTE 567 thermal conductivity 558 material 460 Monolithicity loss of 469 thick-walled structures 456 Monomers BMI 100 ratios 79 reactive 34, 37 Montana talc 248-9, 248 Morphology, resins 505 MPDA/MDA, viscosity 56
Index 1097 Mullite 313,314,318 heat treatment 310 Multi-fastener joints 619-24, 620, 621,622,623 Multi-rowed, fasteners 620 Multifilament, continuous 28 Multilayer locking, continuous fiber 338 Multiple gates 444 ply angles 19-20 stage drawing, PET 224 Multiwarp, weaving 406 Musical instruments 946 Mutagenicity 824 MWK 3-D 409,410 LIBA system 411 MY 0510 (Ciba-Geigy) 53 MY 720 (Ciba-Geigy) 53 Mylar film sheet 574 template 487 N, N, N , N'- tetraglycidyl methylenedianiline 53 N/Cratio 182 Nadic methyl anhydride (NMA) 57 Nanoporous, carbon fiber 183 NASA standards 1063,1064 NASP 972,972,974 National Aerospace Plane, see NASP Natural composites cartilage 958 cortical bones 958 dentin 958 wood 958 Nd-YAG applications 605 solid state laser 605 NDE 733,838-55 blown core 854 bondline adhesive voids 854 condensed core 854 core 854 damage inspection 854 deep delamination 854 degradation 838 disbonds 854 fiber breaks 838,854 foam adhesive voids 854 foreign material 838, 854 inhomogeneities 838 matrix cracks 854 porosity 854 records 733 shallow delamination 854 test methods 733-4734 water intrusion 854
Neat polymers, shrinkage 541 Nepheline 2434 chemical resistance 2 4 3 4 oil absorption 244 particulate fillers 2434 refractive index 243-4 syenite 244 Neutral hydrolysis 887 Neutron radiography 847-9,848 field inspections 848 military aircraft 849 New York, talc 248-9, 248 Nextel 309,309,310,316,403 312 fiber 159 Nibbling 602 Nicalon 310-11, 311,316, 318 Nickel 249 coating 292 electroformed, CTE 589 powder 250 vapor-deposited coating 819 Nitride, coating 296 Nitrogen release, and tensile strength 182 Nomex 2034,269 compressive strength 272,273 cores 270 decomposition temperature 206 density 206 elongation at break 209 filament 206 initial tensile modulus 209 linear thermal expansion coefficient 206 long-term use temperature 206 melting temperature 206 moisture content 206 plate shear modulus 272,273 strength 272,273 specific gravity 209 specific heat 206 tensile strength 209 thermal conductivity 206 Nominal thickness, reinforced composites 390,391 Non-aerospace, applications 935 Non-functional, fillers 382 Non-metallic core removal 862 Non-symmetric bending, testing 783 deformation, cylindrical shells 754-5 laminates 752 Noncomposites 166 Nondestructive evaluation, see NDE testing, and inspection 1059
Nonlinear, bending 747 Nonrecyclables disposal of 899-900 incineration 899-900 landfill 899-900 Nonwoven fabrics 402,402 textiles 418-20 3-D technology 418 design methodology 419 orthogonal 3-D 418,419,420 processing 418 structural geometry 419 Normalization 721 compression strength 721 Notation, stacking sequence 619 Notched fatigue behavior 802 Novolac resin homopolymer property 102 structure precursor 102 supplier 102 tradename 102 Novoltex 418,418 NR-150 char yield 82 chemistry of 77,78 density 82 elongation 82 flex modulus 82 flex strength 82 fracture toughness 82 neatcured 82 Rockwell hardness 82 tensile strength 82 Tg 82 thermal expansion 82 Nuclear magnetic resonance (NMR) 982 Nylon 6/6 PCI-glass elongation 899 flexural modulus 899 flexural strength 899 glass content 899 izod impact 899 mold shrinkage 899 tensile strength 899 Nylon modulus and temperature 210, 211 peel plies 673 imprint 682 tensile strength and temperature 210,211 weld line strength 544 O-rings 443,953 nitrile rubber 443 Observed life, AS-4/PEEK(APC) 806,807
1098 Index Octyl, decyl glycidyl ether blend, viscosity 62 Off-axis tension, testing 782 Oil absorption nepheline 243 number, particulate fillers 251 containment, Spectra 2 3 3 4 and gas applications 946-7,947 industry, pultrusion 521 platforms 9245,925 One-dimensional consolidation equation 584 flow, compression molding 583-4,584 stress analysis 627 One-sided, pulse-echo testing 842 Onion-skin, microstructures 177 Opaque, aramid fibers 221 Open core evacuation, aircraft 865 Open hole coupon strength, joints 626 Open packing 403,403 Open unloaded holes 614 Openings 373,373 Operational requirements 737 Optical properties aramid fibers 215 fiberglass 235 Organic solvents, thinning 65 Orientation angle, yams 412 Orifice plates 495 Orthogonal nonwoven fabrics 163, 164 Orthopedic applications 95943 FRP 959 hydroxyapatite (HA) 958 PMMA 962 Orthophthalics 35,39 polyester, blistering 918 Orthotropic layers 742 nonzero stiffness 742,743 thermal coefficients 742 plates, critical load 751 Outgassing measuring 817 and vacuum 817 Oven/press cure, molding 366 Ovens, with thermocouples 429 Overheat, environmental exposure 860-1 Overshoots 531 Oxidation PAN-based fibers 818 resistance 818 carbon-carbon composites 333
4,4'-oxydialinine (ODA) 75,76 P75S/cyanate, moisture vs. time 1010 PlOO tape mechanical properties 728 strengths 728 p-t-butyl phenyl glycidyl ether, viscosity 62 PA6,6, water absorption 127 Packages style of 493-4,493 weight 494 wound 493 Packing fraction 531-2 interfiber 403 powder processing 313 pressure 539 Paint surfaces, inspection 839 PAN 25,169,171-3,172 anisotropicity 185 cyclization 172 cyclized 179,181 dryspinning 175 large diameter 180,181 modified polymers 175 molecule structure 172 spinning 175 stabilization 172,177-80,178,179 stabilized, thermal degradation 181 wet spinning 175 PAN-based fibers density 270 elastic modulus 270 fracture elongation 270 oxidation 818 tensile strength 169, 270 PAN-based tow manufacturers 298 trade names 298 Para-aramids available 203 creep 212 fatigue 213 stress rupture 212 Para-phenylene terephthalamide 26 see also aramids Parallel axis theorem 694 Part consolidation 386 ejection 385 geometry 526,530 inspection 733 methods used 733-4 removal 446 hydraulic ejection 446 system 541
shape, and tooling 589-94 slippage, during cure 432 Particle accelerators 976-9,977,978 size clay 244 distribution, rotational molding 552 Particulate fillers abrasion 252 alumina trihydrate 245 aluminum oxide 250 antimony oxide 244-5 calcium carbonate 243,247 carbon black 245 clay 244 common 243-6 cost 242 definition 242 electrically conductive 250 elongation at failure 252-3 enduses 253 feldspar 243-4 fire retardants 249 flakes 248 orientation 252 graphite 245 hardness 251 health hazards 252 high density 249 hardness 249-50 thermal conductivity 250 hot spots 242 injection molding 242 kaolin 244 lightweight fillers 249 low density 243 low friction 250 magnesium oxide 250 mica 245-6 microspheres 246 natural 246 nepheline 243-4 oil absorption number 251 organic 243 packing 252 production 2464,247 shrinkage 242,252 silica 244 specific 243 surface properties 250-1 treatments 250 synthetic 246 talc 246 thermal conductivity 242 toxicity 252 types of 242-3
Index 1099 use of 242 Particulates 307 flow, cf. fiber flow 530 inhaling 827 irregular shaped 249 loading levels 528 radiation 816 Parts, RTM process 762 Pastes, adhesives 275-6 Patterns 362-3 control 463-4 PBI 237,237 chemical resistance 237 chemical warfare applications 237 solvent resistance 237 structural formula 237 PBIA 203,203,205 PBLA-based fibers, SVM 210 PBO 26-7 elongation at break 235 fiber 236,237-8,237 specific gravity 235 structural formula 237 tensile modulus 235 tensile strength 235 PDI fibers, sources of information 236 Pedestrian bridges 991-2,992,993 PEEK 31,120 glass transition temperature 812 Peel plies 668, 669 bonded joints 669 dry 672 evaluation 670 imprint 678 nylon 682 sanding 680 nylon 673 polyester 681,681 preimpregnated 672 removing 683 silicon transfer 670 surface preparation 683,870 strengths adhesives 274-5 GFRP 1006 stress 629,631,637-8,637,638, 647,655,662 adhesives 628 eliminating 644-5 lap joints 6374,637,638 linear variation 643 maximum bond 642 reducing 644-5 tapering 646 PEKK 31 PEL 825,826
Penetration 814 Percolation 547 matrix 548 Perlite 249 low cost 249 Permeability 399400,1000 axial 579 bleeder 581 composite layer 581 and fiber volume fraction 399400,400 preforms 438,439,449,578-9 Permeation barrier 918 Permissible exposure limits, see PEL Peroxides 382 Personal protective equipment 828 PET 35,120 ability to float 225 abrasion resistance 224,226,233 anisotropy 224 applications 230,233-4 artificial ligaments 234 availability 231 axial orientation 224 chain folding 224 characteristic temperatures 122 chemical resistance 224,230 cleaning 230 creep resistance 227 crystallinity 223,224 degradation 889 design 232-3 dielectric constant 224 durability 233 dyeing 233 energy to break 227 exposure to moisture 227 fatigue resistance 230 fibers, extended chain 202 filament 225 filter cloths 233 forms 232 gel-spun 225-6 handling 233 high strength 233 hybridization 233 hydrophobic 230 inert 230 machining 233 melt extrusion 224 melting point 225 moisture absorption 224 multiple stage drawing 224 non-conductive 231 non-woven 234 processing temperatures 233 recycling 883 for reinforcing 234 resin penetration 233
scrap 36 seawater resistance 233 shear modulus 120,120 sources of information 232 specific gravity 224 modulus 224 strength 224 stiffness-to-weight ratio 223 strength retention 227 strength-to-weight ratio 223 structure 224,224 surface treatments 231-2 sutures 234 use temperature 225
uv
resistance 224,230,233 stability 233 Phase locked loop, holography 852 Phenolics 504-5 adhesives 255 compressive strength 258 density 258 health effects 834 maximum service temperature 258 polypropylene, density 258 pulling 5 0 4 5 pultrusion grades 505 resins 268,504 resistance to fire 505 shear modulus 259 shear strength 259 SMC 382 spa11 liners 914 tensile strength 258 thermal conductivity 259 Phenoxides allyl functional 101 propenyl functional 101 Phenyl glycidyl ether (PGE), viscosity 62 Phosphoric acid anodized, aluminum foil 820 Phthalic anhydride 35 chemical resistance 35 marine applications 916 melting point 57 thermal stability 35 Physical properties aramid fibers 2057,206 graphite/epoxies 4 Pick-up trucks 936-7 Pigments 516 inorganic 516 titanium oxide 516 zinc sulfide 516 Pins,FRP 959 Pipelines 466
1100 Index rehabilitation 948 Piping 936 systems 926 Pitch 1757,176 fibers 169,352 graphitized 183 flow through spinneret 177 forming coke 173 mesophase 176 precursors 173 carbon yield 197 treatments 173-5,174 pretreatments 173,197 softening 197 spinning conditions 1757,176 stabilization 177 stirring 277 Pitch-based fibers 170 density 270 elastic modulus 2 70 fracture elongation 270 tensile strength 2 70 Pitch-based tow manufacturers 199 trade names 199 Placement, on mold 384-5 Plain strain 801 weave 140,145,405,406 biaxial 399 fiber volume fraction 408,408 triaxial 399 Planar flaws 840 Plane source 453 stress, stiffness 689 Plasma assisted CVD (PACVD) 316 Plasma spraying 300 Plaster masters 5644, 564 follow board method 565,565 NC machining 566 sweep method 565-6 template method 564-5 Plasters, breakout/washout 465 Plastic faced plaster 567,573 preparing 573 Plastic zones 654 Plasticating 538 extrusion 534,535 injection molding 538, 538 Plasticization 107 moisture 112 Plastics, recycling 883 Plate elements 621 reinforced, carbon/aluminum 304 shear modulus honeycomb 2665,266-7
Nomex 272,273 shear strength honeycomb 2665,266-7 Nomex 272,273 stiffness 622 thickness 790 structural hierarchy 778 Platelets 307 Platens, heated 499 Platforms 989 fiberglass 989 well bay 989,990 Pleating 547 Plied-yarn compressive strength 252 count 252 flexural strength 252 modulus of elasticity 252 tensile strength 152 warp 252 weave 252 Plugs 362-3 Ply alignment 689 angle 699-700 boundaries 481-2,481 locating, templates 574-5 orientation 9,lO-11,374 interspersing 19 properties 8,737 reinforcement 730,730 sequence, table 732 shape 482-2,481,482 generating 481 stacking sequence 692 strength 745 Plycosite 255 PMC decontamination 863 evacuation 863 PMDA/ODA polyimide 76 PMMA fatigue resistance 959 orthopedic applications 962 PMR-15 80 air aging 89 chemistry 79-80,79 compressive strength 82 density 82 elongation 82 fracture toughness 82 laminates 88,89 moisture absorption 82 neatcured 82 processing 84 tensile modulus 82 tensile strength 82 T, 82 thermal expansion 82
PMR-I1 cf. PMR-25 90, 93 chemistry of 80-2,80 processing conditions 84-5 PMR-11-50 flex strength 92 PMR-11-30 90 reinforced 90 Poisson’s ratio 19,698 ceramic composites 312 IM6/epoxy 699 M401/F854 699 Polar backscatter, ultrasonic 843 Polar winding 461-2,462 Poly (p-phenylene benzobisoxazole) see PBO Poly (p-phenylene benzobisthiazole) see PBT Poly-2,2’-rn-phenylene-5,5’benzimidazole, see PBI Poly-rn-phenylene isophthalamids 203,203 Poly-p-phenylene terephthalamide, see PPTA Poly-p-phenylene-benimidazoleterephthalamide, see PBIA Polyacrylonitrile, see PAN Polyaluminoxane 159 Polyamide 6,6, characteristic temperatures 221 Polyamide 12, characteristic temperatures 222 Polyamide-imides chemical name 222 processing temperature 222 structure 222 suppliers 222 T 222 l-: 122 trade name 221 Polyamides characteristic temperatures 222 health effects 832, 833 Polyaminoamides, health effects 832,833 Polyarylene ether chemical name 222 processing temperature 222 structure 222 supplier 222 T 122 221 trade name 121 Polybuty lene-terephthalate, characteristic temperatures 122 Polycarbonate, weld line strength 544 Polycrystalline diamond, milling
r”,
Index 1101 596 Polyester 28,34-9 characteristic temperatures 722 chemical name 222 curerate 509 density 505 elongation at break 505 flexural modulus 505 flexural strength 505 glass fiber reinforced 28 heat distortion 505 linear 34,34 modulus and temperature 210, 211 peel plies 681,681 polyarylate 235 processing temperature 222 reinforced 4.42 resins, shut down 515 shrinkage 507 structure 222 supplier 121 tensile modulus 505 tensile strength 505 and temperature 210,211 T, 721 T, 721 trade name 222 unsaturated 35 Polyether ether ketone, see PEEK Polyether ketone ketone, see PEKK Polyetherimide, flexural modulus 545,546 Polyethylene recycling 883 terephthalate, see PET Polyglycidyl ether of o-cresolformaldehyde Novolac melting point 52 viscosity 52 Polyglycidyl ether of phenolformaldehyde Novolac viscosity 52 Polyimide addition chemistry 78-81 aromatic 75 binder solutions 76 chemical name 127 compression molding 545 condensation 78 epoxy resins 6 precursor solutions, see binder solutions processing temperature 227 structure 227 supplier 721 T 727 T: 222 tradename 222
Polymers composites 378 properties 3854,386 resin 382 compound grades 525 crystalline 118, 120 economics 525 extrudate swell, injection molding 540 fire retardant 42-5 frictional coefficients, injection molding 540 inclusion of particulates 528 major processes 526 matched metal dies 546 matrix composites seePMC grinding 600 matrix lamina, fiber fracture 799 melt compressibility, injection molding 540 melt fracture potential, injection molding 540 polymerization condensation 7 5 4 shrinkage at 515 pressure dependent shrinkage, injection molding 540 processes 529,532 unit costs532 processing 525 properties, blow molding 550 rheological concerns 525 selection 525 semi-crystalline 534 processing temperature 120 shear sensitivity, injection molding 540 thermal properties, injection molding 540 thermoplastic, adducts in 526 thermoset 34 viscosity, injection molding 540 Polymethylmethacrylate, see PMMA Polyolefin 549 characteristic temperatures 222 Polyorthoesters 957 Polyoxymethylene cupping 541 shrinkage 541 warping 541 Polyphenylene sulfide, see PPS Polypropylene see also PP and acrylic acid 251 characteristic temperatures 721 mica-reinforced 253 phenolics 258 shear modulus 259
shear strength 259 thermal conductivity 259 viscosity 530, 531 weld line strength 544 Polysulfone see also PSU chemical name 222 flexural modulus 546 processing temperature 222 structure 222 supplier 727 T, 122 Tm 222 trade name 722 weld line strength 544 Polytetrafluoroethylene, see PTFE Polyurethane coatings 813 compressive strength 258 density 258 health effects 834 maximum service temperature 258 shear modulus 259 shear strength 259 tensile strength 258 thermal conductivity 259 Polyvinyl chloride compressive strength 258 density 258 maximum service temperature 258 shear modulus 259 shear strength 259 tensile strength 258 thermal conductivity 259 Pontoon pier 995 Poor bonding, adhesive joints 611 Porosity 90,300,630-1 bond layer 656 NDE 854 preform 439 X-ray imaging 84.4 Ports and harbors 949 Post-curing 84-5, 385,446-7,589 Post-fabrication 732-3 Post-processing 446-7 Post-tensioning 986,987 anchors 987 Pot life 66, 510, 510 anhydride-cured system 72 NMA 66 resins 440,513 Potassium aluminosilicate 2454 metaphosphate 957 Powder alloys 300 consolidation of 343
1102 Index impregnation 228 processing ceramic composites 312-14, 313,314 packing 313 sintering 343 Powdered amorphous thermoplastics (TPs) 105 Power poles 948-9,948 shears 426 trowels 461 Powertrain applications 911-12,911, 923 PP and granulated SMC elongation 897 flexural modulus 897 flexural stress 897 izod impact 897 tensile modulus 897 tensile stress 897 PP / PCI-glass glass content 898 strain at break 898 tensile modulus 898 tensile strength 898 yield stress at break 898 PPD 81 PPS 31 weld line strength 544 PITA 203,205 chemical stability 213 hydrogen bonding 204 structural formula 203 Prebaking 871 Precarbonization 180 Preceramic polymer chemistry 157 Precured patching 867-8 limitations 867-8 Precursors 335 fiber, pyrolysis 157 materials 173 mesophase pitch 335,336 PAN 335,336 pitch 173 polymeric 157 rayon 335 treatments, pitch 173-5,174 Predicted life, AS-4/PEEK(APC) 806,807 Predicting durability 796 joint strength 615 peak stresses, joints 613 Preform advantages 438 ceramic 314 corners 438 designing 438
disadvantages 438 drapability 438 edge definition 438 elastic constant 449 engineering parameters 422 fabricated 438 fiber distribution 438 fiber orientation 422 heating 458 infiltration 299 permeability 438,439,449,578-9, 578 porosity 439 prefabricated 438 processing parameters 421 R T M 436-9,437,438 textile 401-2 thickness 438-9,438 uncured 456 winding 458 Preforming 397422,448 3-D 398 formability 398-9 Preliminary design 710-11,715-23 definition 710 materials 715-22 support data 711 materials, aircraft stabilizer 722 Preload relaxation 729 Prepegging 6-7,7,32-3 advantages of 7 hot melt 7 roving 7 thermoset matrices 7 Prepolymers 105 Prepregs aerospace industry 885 availability 82 braiding 418 co-curing 868,870 combining 117 composite tools 566,567-71 consolidation 117 definition 425 dry 425,430 facing material 255 lay-up 576 low tack 425 over metal substrates 870 ply kits 374 repairs 868-70, 869 resins 352 matrix 577 reuse 884-5 shredding 885 suppliers 109 tapes 430 thermoplastics 357
tooling 562 unidirectional 118 waste 885 hazardous 885 wet 355 Pressing temperature, SiC/Al,O, 325 time, SiC/A403 325 Pressure compliant ring, testing 782 consolidation 577 damage 874,875,876 gradient, infiltration 338 hulls, marine applications 922 pulsation, infiltration 338 vessel, acoustic 472,472 vessels 471 X-ray backscattering 846 PRESTEK 992 Prestressed concrete 985-6 anchorage 985-6 bridge deck 986,986 bridges 987 steel tendons 985 Preventive maintenance, composites 880 Pricing aramid 216,716 Dyneema 232 graphite fiber 716 Hoechst Celanese 232 Spectra 232 Primary Adhesively Bonded Structures Technology (PABST) 6674,669 Principle strengths 803 Processing 64-9 composites 879 glass fibers 24 parameters, preforms 422 polymers 525 speeds cf. pull loads 506,506 pultrusion 514 technology, braiding 413-15 temperature PET 233 polyamide-imides 221 polyarylene ether 222 polyester 222 polyimides 222 polysulfones 122 thermoplastics 115,52552 time, terephthalic resins 36 window 176 Procurement costs, materials 722 Producing, carbon/aluminum 303 Production costs, aircraft 1024
Index 1103 MMC 291 particulate fillers 2464,247 Profile extrusion 532 Programmed winding 469-70 Propeller shafts 911 Properties aramid fibers 205-15 boron 161 casting, CE 108 ceramic fibers 161 elastic 194-5 highsilica 161 Kevlar 205 lamina material 760 polymer composites 385-6,386 quartz 161 Technora 205 thermoplastics 115 Twaron 205 using recyclate 892 Vectran 235-6 Proprietary specifications 1066-7 Propulsion shafting 925 systems 1041,1041 Prostheses 963 Protective coatings 292 composites 879 Protruding head, fasteners 624,624 Pseudoisotropic laminates, CMEs 1011
PTFE 362 see also Teflon compression molding 545 velocity cf. temperature 843-4, 844 Pull forming 490-2,490 automation 492 cf. pultrusion 490,491 curved 490,491,491 straight 490,491-2 loads 512 cf. processing speed 506,506 epoxyresins 510 winding 489,490,496-7 Pullers continuous belt 501 intermittent 501 Pulling 501-2,501 continuous 501 Pulp molding, short fibers 337,338 Pulse-echo testing advantages 842-3 c-scan 842,842 one-sided 842 ultrasonic 842-3,842 Pultrusion 17-18,33,117,488-521, 529,532
additives 515-18 advantages 489 applications 518-21 basic process 489 cf. extrusion 488-9 cf. pull forming 490,491 civil engineering 519-20 and compression molding 490 construction 519-20 continuous production 489 creel 492-5 w e d 489-90 definition 488 dies 497 electrical equipment 518-19 epoxy 511 fibers used 503 fillers 516 FRF 47,517 hammer handles 491 history 488 injection 497,498 labour requirements 489 machine 488 oil and gas industry 521 pigments 516 processing speeds 514 products 489 pull-winding 496-7 resins 504 impregnation 495 shapes 989 shut down 515 sloughing 514 sporting goods 1045-6 start-up 514-15 surface finish 514 thermoplastics 116 three zone model 505,505 transportation 520-1 tubular structures 496-7 vertical 495-6 wastage 489 and winding 496-7 windmills 521 Pure bearing load, joints 626 Purging 515 Putties, adhesives 275-6 PVC foam 257 crosslinked 917 linear 917 Pyrex 318 Pyrolysis 159,182,888 carbon fibers 335 carbon-carbon composites 341, 342 carbonized organic composites 340
definition 888 precursor fiber 157,159 recycling 896-8 tooling 341 yieldsfrom 895 Pyromellitic dianhydride (PMDA) 75,76 Quadratic failure 695-7 criteria 707-8 Qualification tests 1060 Quality controls, composites 880 Quartz applications 163,165,166 drawing 157 fibers 24-5 fused 163 forms 161-3 fusedrods 157 manufacture 156-7 properties 161 Quasi-isotropic laminates 9-11,697 materials 687 modulus 12 Quaternary ammonium salts 382 R-glss 134 Racetracking 443 pattern 461-2 winder 462 Racing yachts 353-4 Rackets 1049-50,1050 braiding 1050 compression molding 1049 filament winding 1050 injection molding 1050 resin transfer molding 1050 Radar signals, transmission 472 transparency, foam 256 Radial microstructures folded 177 withwedge 177 Radiation 3,816 Bremsstrahlung 816 ionizing 816 particulate 816 solar 816 thermal 816 ultraviolet 817,817 Radio frequency see also RF heating 499 Radius of curvature 459 sheets 186,187 Radius of twist 459
1104 Index Radomes 1024 composites, quartz reinforced 110 Rail shear, testing 783 Railroad applications 912-14 rolling stock 949-50 Railways 950 Random, microstructures 177 Rapid cure, resins 440 Rate of relaxation, water absorption 811 Ratio, glass-to-resin 1504,153 Raw materials, falling cost 838 Rayon 25 RD-1 (Ciba-Geigy) 61 RD-2 (Ciba-Geigy) 62 Reaction efficiency, atomic oxygen 814 Reactive rubber, tougheners 106 Reactivity, resins 440 Reaming 598 Rebars 984-5,984 fiberglass 999 placement 985 pultruded 999 Recreation, design requirements 713 Rectangular plates 749-53,749,751 in-pane loading 750-1,751 Recyclability,MMCs 944 Recycled materials 889-99 NBC 897 PET 894 phenolic 894 PP 895,896 SMC 890,892,893 Recycling see also reuse, disposal automotive industry 901-2 BMC 890-3 building construction 901-2 electrical parts 902 environmental aspects 902-3 Europe 884 Germany 887,902 glass-filled PP 893 granulation 886-7 heat cleaning 889 history 883-4 Japan 887,888,903 legal aspects 902-3 mechanical 886-7 milling 886-7 organizations 903 PET 883 phenolic composites 893 plastics 883 polyethylene 883
polyurethane composites 893 pyrolysis 896-8 reinforced composites 388 SMC 890-3 Sweden 887 thermoplastics 893-5 Redux 255 bonding 667,669 Re-entry temperatures, spacecraft 816 References E-glass 774-7 graphite epoxies 774-7 Kevlar 774-7 lamina property 772-3,774-7 S2-glass 774-7 Spectra 774-7 Reflectors 1015 ACTS 1017 Refractive index feldspar 243-4 Kevlar 215 nepheline 243-4 Regrind RIM density 895 elongation 895 flexural modulus 895 hardness 895 tensile modulus 895 tensile strength 895 Rehabilitation bridges 988 building construction 988 chimneys 988 concrete 988 steel 988 wood 988 Reinforced composites 712,713 applications 38carbon fiber 358-9 chopped-fiber 355 continuous-fiber 355 continuously 358 corrosion resistance 387 corrugated configuration 389 dimensional stability 387 edge stiffening 390,391 edgetuming 390 electrical resistance 387 high strength 387 inserts 392 light weight 387 minor impact resistance 387, 388 molded-in color 388 nominal thickness 390,391 non-structural requirements 388
recycling 388 ribs 389,391,391 shell and plate construction 389,389 size and shape 388 spray-up 355 structural requirements 388 surface quality 387 concrete 983-8 beams 1000 cracking 1000 E-glass 984 steel reinforced 983 epoxies 442 polyester 442 reaction injection molding, see RRIM Reinforcement CE 111 circumferential 461 composites continuous unidirectional 318-20,318,319 discontinuous 320-1,320 configurations 461-4 continuous 308-11,309,310,502, 762 definition 378 E-glass 383,916 geometry 797 glass fibers 378,383 graphite 353 heat cleaned 898-9 helical winding 461 hoop winding 461 Kevlar 383 knitted 916-17 local 730,730 marine applications 916-17 materials 832, 835, 835 RTM 435 monofilament 311,311 multi-directional 718 plies 730,730 polar winding 461-2,462 S2-glass 383 spatial 471 specific heat 533 thermoplastic polymers 526 three-dimensional 456,801 two-dimensional 456 types fibrous 795,796 unidirectional 438 volume 510-11,510 woven 361,916-17 Reinforcing bars, see also rebars ceramics 307-11,308,308,309,
Index 1105 310,311 load carrying capabilities 307 Relaxation, environmental effects 1000 Release agents 362 Releases 429-30 fluorocarbon 361,362 molds 361,362 silicon 49-30 Remaining life 804-7,804,805,806,807 strength 797,804-7,804,805,806, 807 estimating 797 loading history 805 normalized 805,805 predictions 806 Removal of peel ply 631 Repairs adhesives 869 advanced fibers 877 aircraft 857-80 autoclave 868 heat application 874-7 heating blankets 869 honeycomb panels 871-7 joint preparation 871 mechanically fastened 877 metal bondments 871-7 methods 880 non-autoclave 868-9 prepreg 868-70,869 resins 869 skill requirements 878 standardization 878 surface preparation 870-1 technical training 878 vacuum pressure 869 Repeatability 8 turning 603 Representative volume 796,804,804 discontinuous 804 Reproducibility 3 Reproductive toxicity 824 Residual strength, AS-4/PEEK(APC) 806, 807 stresses, control of 470,472 Resins 382 bathlife 496 bleed sequence 367 catalyzed 507 changing 497 chemistry 99-1,382 comparing 505-7,505,506 consumption 513 content 425 byvolume 503 fibreglass epoxies 153
marine laminates 920 minimizing 673 epoxies 382,504 and fibers 425 flow 578-80 consolidation 576 continuity condition 582-3 Darcy’s law 578 gelation 499 high performance 122 high temperature 818 impregnation 495 ingredients 513 injection 117 molding 122 system 433,434,440-1 thermoplastics 226 isophthalicpolyester 504 matrix 504 prepreg 577 mixing 513-14 temperature 441 viscosity 441 moisture absorption 440 morphology 505 penetration, PET 233 phenolic 504 polymerization 499 pot life 440, 496,513 pultrusion 504 rapidcure 440 reactivity 440 recycling 515 removing excess 576,673 repairs 869 replenishing 513 safety 831-2 selection 45,440,504 solvent based 431 Tgpoint 440 thermoplastic matrix 496 mechanical properties 122 toughness 440 transfer molding 17,374,492 see also RTh4 marine applications 919-20 rackets 1050 transverse modulus 1056 tensile strength 1056 uncatalyzed 506 unsaturated polyesters 382 used RTM 440 selecting 440 vinyl esters 382,504
viscosity 64,440,578,579,579 and cure cycle 354,355 volume fraction 478 weather resistant 813 Young’s modulus 440 Resistance chemical 39 to fire,phenolics 505 water 39 Resonance, ultrasonic 843 Respirators, filter 830 Respiratory protection 830 Restoration of coatings 867 Resulting strain state, laminates 693-4 Reticulating films, adhesives 276 Reusable bags, silicone rubber 371, 371 Reuse see also recycling, disposal appliances 901 automobiles 900-1 carbon fiber 884 cutting 885 and disposal 883-904 glass fiber 884 hybrid composites 883 prepregs 884-5 sheet molding compound 883 shredding 885 technologies 885-9 RF preheating 513,513,515 Rheology 527-31 definition 527 Ribs designing 391 geometry 391 moldmaking 390 parameters 743 reinforced composites 391,391 Rigid adherends 633,633,634 fiber 802 tool, thickness 450,450 Rigidity, and flow 67 Ring specimens bending 785 testing 782,784 Rings compression testing 782,788-9 hoopwound 789 thick-walled 789 thin-walled 789 Risk 823 definition 823 Rivets, aluminum 668 RK, properties of 270 Rocket motor
1106 Index cases 23,465 X-ray backscattering 846 design requirements 713 nozzles, carbon-carbon 712 Rockwell hardness, NR-150 82 Rods, testing 784 Roll forming 117,289 wrap process, sporting goods 1045 Roller press 426,427 Rollers 461,495 Rolling carbon/aluminum 303 compaction 477 rate of 302 solid phase 302 solid stage production 291 table 427,427 in vacuum 300-6,301,302,305 temperature conditions 300 Rolls and air shafts 950-1,951 Rotational molding 529,532,550, 551 crosslinking 551 crystallization kinetics 551 molecular characteristics 552 particle size distribution 552 polymer properties 552 thermal properties 551 zero-shear viscosity 551 Router speeds 289 Routing 602 Rovings 163,164 carbon fiber 502 for chopping 146 continuous glass 492 fiber 492 fiberglass 136 glass 502 harder 146 nonwoven 164 Sglass based 147 single strand 136,146 woven 136-7,137,164,917 reinforced 237 yields 136 RRIM 907 RTM 433-54'762 advanced technology 447 advantages 434-5 cavity design 441 compounds, rheology 101 curecycle 433 curing 445-6,445 cycletimes 433 disadvantages 434-5 gate and vent 441 glass transition temperature 440
heating/cooling 441 injection pressure 433 low cost parts 433 moistw absorbance 440 mold sealing 441 optimum viscosity 440 reinforcement materials 435 resin 440 T6 446 sporhng goods 1046 tensile elongation 440 tensile modulus 440 vacuum assisted 433,453 viscosity 440 Rubbery sheet deformation 526 Rule of mixtures 150,318,588,697 Runner system 444 S2-glass 134 compression properties 360 constituent properties 766 cost of 917 elastic properties 764-7 references 774-7 specific gravity 383 strength properties 764-7 tensile failure 383 tensile modulus 383 tensile strength 383 thermal expansion 383 S twist, yam 138 Sglass 23,24,425 based, rovings 147 density 503 elongation at break 503 epoxies 252 fibers, stress rupture 212-13,213 filament strength 24 meltingpoint 24 modulus 24 tensile modulus 503 tensile strength 503 SS2-glass reinforcement 383 specific modulus 1049 specific strength 1049 tensile modulus 360 tension-tension ratio 1049 SS2-glass, fiber modulus 360 Striazine ring 99,100 SAAB JAS39 Gripen, composite applications 1034 SACMA standards 1061-2,1062 Safety applications 943-4,944 aramid fibers 222 epoxy resins 831-2,832 factor, formulas 285 hazards, composite processes 831
resins 831-2 Saffil 159 grades of 159-60 Sailboats,design requirements 713 Same materials, sandwich structures 744 S A N ,weld line strength 544 Sand bags 874 Sanding 672,672,678-80,679,683 peel-ply imprint 680 scuff 679 Sandwich adhesive flow 288 beamtype 281 buckling 282 construction 694 aramidfibers 221 conversion to 255 core crushing 282 selection 276,278,279,284 sue 288 deflection 284 limitations 281,284 design 281,284 compressive modulus 281 core 281 facings 281 notation 280-1 designing 276-80,277,278,279 dimpling 282,284 fabrication 276 face wrinkling 282,284 facing failure 282 high temperatures 288 manufacturing 287-9,289,290 modes of failure 282,282 shear crimping 282 skin materials 284 structures 467-8 design 467-8 different thicknesses 744 elements of 254,254 same materials 744 space vehicles 254 thickness 284 transverse shear failure 282 wall 742,743 transverse shear deformation 743 Satellites central cylinder 375-6 design requirements 713 Satin weave 405,406 Saturated acids 37 Sawing 601 Scarf adhesive joints 628,629 joints 611,623,645
Index 1107 shear stress 649 distributions 6474,648, 649 Scarfing 6234 Scientific applications 96740 Screws, injection molding 538 SCRIMP 919 Scuff sanding 679 Sealants, composites 879 Sealing 577 composites 302 Seals 953 Seawater resistance, PET 233 Secondary, drilling 598 Section failure curves 615-16,615 load, last fastener 620 stress 615 Secured Modular Automotive Rail Transport (SMART) 914 Selection process, joining 727,728 Self-adhesiveskins 276 Self-screening, aramid fibers 215 Semifabricatedcomposite, winding 468 Sensitization 824 cross 824,825 equations 825 Sensors pressure 499-500 temperature 499-500 Separation 597 Shallow delamination,NDE 854 Shear 394,395 coupling, coefficient 193 crimping, sandwich 282 deformation 448 cylindrical shells 756 edges 394 failure 195 in-plane 783,789-90 interlaminar 790-2 knifeedge 394 lag analysis 633-9 modulus ABS 259 cellulois acetate 259 epoxies 259 IM6/epoxy 699 M401/F854 699 PET 120 phenolics 259 polypropylene 259 polyurethane 259 polyvinyl chloride 259 skinned molded foams 259 properties aliphatic system 70 aromatic system 71
rail test 789 strength ABS 259 AFR700B/S2 laminates 93 carbon fibers 295 Celion 87 cellulois acetate 259 epoxies 259 interlamina 196 phenolics 259 polypropylene 259 polyurethane 259 polyvinyl chloride 259 skinned molded foams 259 stress 636,636 5-step 649,650 10-step 649,651 adhesive 632-7,633,634,635, 636 distribution 636 scarf joints 6474,648,649 maximum bond 642 peak-to-average 648 scarf joints 649 TEOS 346 test 688 in-plane 196 Shearing, weave 439 Shearout failures 618,619 Sheet extrusion 532 molding 374 molding compound seealso SMC reuse 883 radius of curvature 186,187 Shelf life, before molding 31 Shell and plate construction, reinforced composites 389, 389 Shells, netted-ribbed 473 Shims 570,572 Shipboard armor 925 Shipments 955 Short beam shear properties, Spectra 226,226 fiber mat, flexural strength 401 fibers 337,338 term exposure limit, see STEL Shredding mobile 887 prepregs 885 Shrink factors 557-8 Shrinkage 5,179,506,507-9,507, 508 fillers 508 following gelation 508 neat polymers 541
particulate fillers 242,252 polyester 507 polyoxymethylene 541 rate 509 recycled SMC 892,893 volumetric 507 Shut down, pultrusion 515 Sialon 329 SiC/AJO, density 315 pressing temperature 315 time 315 SiC/SiC composites 401,401 Signal-to-noise ratio, ultrasonic 834 Silanes 250-1 coupling agent 147 organofunctional 147 Silar fracture strength 325 toughness 325 work of 325 Young’s modulus 325 Silica crystalline 249 fillers 527 flocculated varieties 244 fumed 247-8 fused 244 natural 244 particulate fillers 244 surfaces 244 thixotropic effect 244 Silicates, fillers 527 Silicon carbide 27,319 coating 294 CVD 157 fillers 527 wheels, grinding 600 nitride, whiskers 161 releases 429-30 Silicone rubber 365,366 CTE 589 reusable bags 371,371 Silver flake 250 Simple twist 138 Single head, p d winding 496-7 lap joints 626,627,628,729 bending deflections 641 load path eccentricity 637, 637 shear, joints 624 strand, rovings 146 strap, adhesive joints 628 Yam
1108 Index compressive strength 151 count 151 filling 151 flexural strength 151 modulus of elasticity 151 tensile strength 251 warp 151 weave 151 Size, meteoroids 815 Siziig 802 agent 495 fiber reinforcement 435 Skill requirements, repairs 878 Skin core, bonding 917-18 creams 829-30,830 grafts 963 materials, sandwich 284 penetration aircraft 858,859 foil tape repairs 858 protection 828-30 Skinned molded foam compressive strength 258 density 258 maximum service temperature 258 shear modulus 259 shear strength 259 tensile strength 258 thermal conductivity 259 Skybond 76-7,77 dielectric strength 95 electrical properties 95 elongation 86 flex modulus 86 flexstrength 86 insulation resistance 95 processing conditions 84 surface resistivity 95 tensile strength 86 volume resistivity 95 weight loss 86 Skylab oxygen tank 471,471 slip forming 547,547 Slot camera, X-ray backscattering 846,846 Sloughing 508,510 SlUrry-Spinning 159 Small boats 921 SMC 380-1,890-3,892,893,895-6, 897 automobile industry 381 currentuse 381 early applications 380 fatigueproperties 386,387 flexural modulus 386,386 flexural strength 386,386 formulation 384
grill opening panels 907 IZOD impact 386,386 lower density 381 manufacturing 381 molding 384 phenolics 382 recycling 890-3 regrinding 890,891 specific gravity expansion 386, 386 tensile modulus 386,386 tensile strength 386,386 thermal expansion 386,386 SMEs 952 Snowflake P.E. 247,247 Sodium-calcium-aluminumpolyphosphate 957 Soft gauge, machine tools 364 Soft tissues, composites for 963-4 Sol-gel technology 156-7 Solar absorptance 816 array backing 820 panels 353 radiation 816 Soldering, carbon/aluminum 303 Solid conveying, extrusion 534,535 phase, rolling 302 rubber, tooling 670 stage production 291 state laser, Nd-YAG 605 Solution sedimentation,MMC 291 Solvents 812,8357,836 absorption, anhydride-cured system 72 aramidfibers 204 impregnation 117,118 polar 175 removal 159,175 resistance PBI 237 thermoplastics 115,127 toxicity 361,362 trichloroethylene 64 Sonar domes, marine applications 922 Sonotrode material 607,608 non-rotating steel 608 Sources of information Armos 221 Kevlar 221 PBOfibers 236 PDIfibers 236 SVM 221 Techora 221 Twaron 221
Vectran 236 Sources of pressure 460 Space 813-17 atomic oxygen 813-14 composite systems 814 debris impacts 81415,814,815 meteoroid impacts 81415,814, 815 shuttle 344,345,347,472 tooling development 1015-16 Spacecraft 968-72,969,970,971 heat sources 816 LDEF 970-2,970,971,972 re-entry temperatures 816 thermal cycling 818 truss structure 968-70 Spa11 liners, phenolic 914 Spallation 814 Sparging 1745 gas 174-5 Spatial reinforcement 471 Spatially sewn structures 456 Special tests 792 Specific gravity Armos 208 Dyneema 225 E-glass 383 expansion 386,386 fiberglass 235 graphite 383 Hoechst Celanese 225 Kevlar 208,383 Nomex 209 PBO 235 PET 224 recycled SMC 890,892,893 S2-glass 383 Spectra 225 SVM 208 Technora 209 Teijinconex 209 Tekmilon 225 Twaron 208 Vectran 235 Specific heat aromatic system 71 definition 532-3 fillers 533 Kevlar 206 Nomex 206 reinforcements 533 Teijinconex 206 Twaron 206 Specific modulus 715,716 aluminum 1049 carbon 1049 definition 715 E-glass 1049 Kevlar 1049
Index 1109 Kulon 305 PET 224 S2-glass 1049 Spectra 1049 steel 1049 titanium 1049 VMN-4 305 Specific stiffness, aramid fibers 207 Specific strength aluminum 1049 aramid fibers 207 carbon 1049 continuous fiber 162 E-glass 1049 Kevlar 1049 metals, vs. composites 1047,1048 PET 224 S2-glass 1049 Spectra 1049 staple 162 steel 1049 titanium 1049 whiskers 162 Specifications materials 722 and standards 1059-67 Specimen geometry, honeycomb 263 Specimen width 780 Spectra 23,26,223,228,229,761 attributes 358 availability 231 axial compressive properties 226, 226 axial tensile properties 226,226 carbon 359 constituent properties 767 creep resistance 227,22&9,230, 233 crystallite melting point 26 density 503 dielectric constant 231 elastic properties 7 6 4 7 elongation at break 225,503 fiber type 225 flexural properties 226,226 graphite, hybrids 354 marine applications 917 for oil containment 233-4 pricing 232 references 7 7 4 7 short beam shear properties 226, 226 specific gravity 225 modulus 1049 strength 1049 strength retention 230 tensile modulus 225,503 tensile strength 225,503
Spectra Shield 234 Spectroscopy fluorescence 69 infrared 69 Raman 69 Speed sensitivity 463 Speed tape repairs, aircraft 859 Spherical shapes, filament winding 466-7 Spherulites 120 Spin process, Technora 213 Spinning 1 7 5 7 conditions, pitch 175-7,176 PAN 175 temperatures 175 Spirit of Australia 354,376-7, 376, 377 carbon/epoxy mast 377 Split rings, testing 784 Sporting goods 1044-52 braiding 1045-6 filament winding 1046,1047 manufacturing techniques 10459 pultrusion 10454 roll wrap process 1045 RTM 1046 Sports cars 910 Spray lay-up, short fibers 337,338 Spray metal, tooling 442 Spray-up 436,437 Spraying ionic 300 plasma 300 Springback 557 GFRP 1013 Springer’s model 578,581,582,585 compacted plies 588 numerical schemes 586 Sprue and runner 541,542 Square packing 403,403 Squeeze flow 548 SRIM 433 machine 434 Stability color 43 PBOfiber 238 and temperature 49 thermal-oxidative 75,92 Stabilization anisotropic pitches 178 commercial 178 oxygen addition 173 PAN 172,177-80,178,179 rate of 180 Stabilizers covers, boron-epoxy 1030-1 uv 43 Stacking sequence
laminate 616,630 notation 619 Stackup, laminate pre-bleed 569 Stainless steel, thermal data 818 Stamping 545 Standard tests, composites 880 Standards, organizations 1060-1 Staple 15940,216 alumina 159 alumina-silica 160 density 162 description 162 diameter 162 elastic modulus 162 green 159 manufacturers 162 specific strength 162 thermal expansion coefficient 162 trade names 162 unfired 159 Static mixers 441 Steel CTE 558 density 170,558 dies 42&7,429,430 thermal expansion 507 elastic modulus 170 fracture elongation 170 linear thermal expansion 705 rehabilitation 988 selection, mold 393 specific modulus 1049 specific strength 1049 tendons, prestressed concrete 985 tensile strength 170 tension-tension ratio 1049 thermal conductivity 558 Steering 482,483 definition 482 radii 482 Stefan-Boltzmann law 850 STEL 826 Step lap joints 628,629,6454,645,649, 650,651,652,659 design 652 Stiffness adherends 636 carbon-carbon composites 333 changes microcracking 796 with time 800 coefficients constitutiveequations 743-4 determining 737 IM6/epoxy 702,703 and coupling 700 cylindrical shells 756 matrix, laminates 693
1110 Index plane stress 689 to weight ratio, PET 223 tooling materials 577 unbalance, adherends 629 weightratio 21 Stitched fabric 436,436 Stitching 801 Storage 357-8 freezer 425 Strains at break, PP/PCI-glass 898 displacment, equations 739 energy, adhesives 653 graphite/epoxies 4 measuring 779 to failure 319 Strand mat continuous 503 fiber structure 503 Strand properties Kulon 306 VMN-4 306 Strands 164 Strap joints, tapering 645 Strength alumina silicate 309 aramid 718 bending 791 anddensity 23 efficiency, carbon-carbon composites 344 envelopes 802,802 evaluation 737 M60J tape 718 PlOO tape 718 properties E-glass 764-7 graphite/epoxies 4,768-71 laminate 764-7 predicting 318,318,319 S2glass 764-7 T50/F584 epoxy 726,726 ratio Ih46/epoxy 701 laminates 696 retention 213 aramids 230 definition 241 Kevlar 205,215 medical applications 957 Spectra 230 Technora 205 structural 795 T300fabric 718 tests, joints 626 three-dimensional 489 tooling materials 577 values carpet plots 719
T50 graphite 724,724 weight ratio 21 aramid fibers 223 PET 223 stress analysis joints 611 one-dimensional 627 concentration 610,613,614,788 distribution, double lap joint 654 interlaminar 738 patterns, laminates 322 residual 466 resultants, laminates 691-3 failure 24 Kevlar 212-13,213 para-aramids 212 Sglass fibers 212-13,213 straincurve 319 thermally induced 640 thermoelastic 466 transfer 149 Stretching elongations 633 Strip heaters 499 Structural applications 709 aircraft 1027-8,1027,1028 formula,PW 237 geometry braiding 415-16 nonwoven textiles 419 hierarchy 778 reaction injection molding see SRlM testing 792 structure aramid fibers 204-5 PET 224,224 polyamide-imides 121 polyarylene ether 121 polyester 121 polyimides 121 polysulfones 121 Styrene addition 34 polyester 382 Submarines 473 Substrates 157 heating 293 Subsurface damage, inspection 839 Succinic anhydride (SA), melting point 58 Superconducting magnetic energy storage systems, see SMEs SuperCollider magnets 978,978 Supennite 247,247 Suppliers
polyamide-imides 721 polyarylene ether 121 polyester 121 polyimides 121 polysulfones 121 prepreg 109 thermoplastics 119 Supported films, adhesives 276 Surface area 607-8 compressive stress 321 crusting 339 of curvature 479 damage 322 defect healing 295 disbonded 677 energies 316,316 finish pultrusion 515 tools 597 flaws 840 geometry concave surface 484 fiber placement 484-6 radii of curvature 484,486 types of 484 ply,veil 503 polish, compression molds 395 preparation 630-1,667-84 abrading 669 adhesive joints 611 aerospace industry 669 aluminum 871,876 automotive industry 669 cuttingplies 870 grit blasting 870-1 history of 667-71 non-autoclave 871-2 peel plies 683,870 removal 631 repairs 870-1 testing 669 properties, particulate fillers 250-1 quality, reinforced composites 387 resistivity fiberglass 135 Skybond 95 roughness, tooling materials 577 slick 682 speeds,grinding 600 treatment 190-1 waviness 603 weave, yams 407,407 Sutures, PET234
SVM availability 220 elongation at break 208
Index 1111 initial tensile modulus 208 PBIA-based fibers 210 sources of information 221 specific gravity 208 tensile strength 208 Sweden, recycling 887 Sweep method, plaster masters 565-6 Swelling stresses, water absorption 811 Symmetric, laminates 692,749 Symmetry 9,10-11 Synergism 43 Syntactic foams, adhesives 275-6 Synthetic marble 253
T50 graphite material properties 724, 724 strength values 724,724 T50/F584 epoxy material properties 726,726 strength properties 726,726 T300 fabric mechanical properties 728 strengths 718 Table rolling 425-32 equipment, suppliers 431 mandrels 429 materials 429-30 Tabs 787 Tack 3567,479 controls 356-7 desired 479 excessive 357 heavy on prepregs 357 requirements 357 Tactical aircraft, material weights 1027 Talc 246 composites 246 impact strength 246 Italian 248-9,248 low cost 246 Montana 248-9,248 New York 248-9,248 polypropylene filled 246 Tank solution processing 871 Tap test 840 laminar type flaws 840 Tape 164 braiding 418 collimated 359 fiberglass 146 non-woven 359 placement 456 reinforcement 25 slit 478 unidirectional 298,359 Tapered
adherends 653 adhesive joints 628 Tapering adherends 627,629,644-51,645, 647,648,649,650,656,658 advantages 646,647 peel stresses 646 strap joints 645,647,648 double 646,647,648 Tappers, mechanical 840 Tateho fracture strength 325 fracture toughness 325 work of fracture 325 Young’s modulus 325 Teglass 134 Technical training, repairs 878 Technology reuse 885-9 winding 458-66 Technora 204 availability 220 breaking strength 214 chemical resistance 224 chemical stability 214 decomposition temperature 206 density 206 elongation at break 209 ether linkages 210 fatigue resistance 213 filament diameter 206 filament shape 206 formsof 216 initial tensile modulus 209 linear thermal expansion coefficient 206 long-term use temperature 206 melting temperature 206 modulus and temperature 210, 211 moisture content 206 properties 205 sources of information 222 specific gravity 209 spin process 213 strength retention 205 tensile strength 209 and temperature 210,211 thermal conductivity 206 Teflon 250,360,834 Teijinconex 205 decompositiontemperature 206 density 206 elongation at break 209 filament diameter 206 filament shape 206 initial tensile modulus 209 linear thermal expansion coefficient 206
long-term use temperature 206 melting temperature 206 moisture content 206 specific gravity 209 specific heat 206 tensile strength 209 thermal conductivity 206 Tekmilon 223,228 availability 231 elongation at break 225 fibertype 225 specific gravity 225 tensile modulus 225 tensile strength 225 Telescopes 1014-15 Temperature 817-19 consolidation 577 control, dies 511-13,512, 513 cryogenic 817 modulus 212 decomposition 76 elevated 817 extremes, GFRP 1004-21 glass transition 75 gradients 470-1 profile, dies 511,512 resistance, aramid fibers 205 spinning 175 stability 49 dies 514 and viscosity 1724,174 Template method, plaster masters 564-5 Templates ply locating 574-5 trim and router 574 Tenacity, definition 241 Tennis rackets 463 Tensile elongation 41,42 recycled PET 894 RTM 440 failure 383 forces, transmission 787 load, fibreglass epoxies 253 modulus aramid fiber 715,716 A S 4 carbon fiber 123 BMC 386,386 carbon fibers 335 carbon (Type T300) 503 carpet plots 719 Celion 87 continuous ceramic fibers 309 Dyneema 225 E-glass 360,383,503,715,7l6 epoxy 522 graphite 360,383,715, !7l6 Hoechst Celanese 225
1112 lndex Kevlar 383,503 Kulon 306 lot-to-lot variation 725 LPMC 386,386 marine laminates 920 PBO 235,238 PMR-15 82 laminates 88 polyester 522 PP and granulated SMC 897 PP/PCI-glass 898 recycled NBC 897 recycled PP 895,896 recycled SMC 890,892,893 r e g r i n d m 895 RTM 440 S2-glass 360,383,503 SMC 386,386 Spectra 225,503 and strength 21,21 Tekmilon 225 thermoplastic resins 222 Vectran 235 vinylester 522 VMN-4 306 ZMC 386,386 properties aliphatic system 70 anhydride-cured system 72 E-glass 360 graphite 360 testing 38 stiffness, aramid fibers 209 strength ABS 258 AFR700B/S2 laminates 93 aluminum 270 aramid fibers 209 Armos 208 AS-4 carbon fiber 123 BMC 386,386 carbon fibers 169 carbon (Type T300) 503 carbon-carbon composites 333 carbon/ aluminum MMC 300 carpet plots 720,720 CE 208 Celion 87 cellulois acetate 258 continuous ceramic fibers 309 Dyneema 225 E-glass 24,383,503 epoxies 258,512 fiberglass 134 epoxies 253 graphite 383 and higher temperatures 182 Hoechst Celanese 225
Kevlar 208,383,503 Kulon 305,306 lot-to-lot variation 725 LPMC 386,386 marine laminates 920 and microtexture 188 and modulus 21,21 and nitrogen release 182 Nomex 209 NR-150 82 nylon 6/6 PCI-glass 899 PAN-based fibers 169,270 PBO 235,238 phenolics 258 pitch-based fibers 270 plied-yam 252 PMR-15 82,88 polyester 511 polyurethane 258 polyvinyl chloride 258 PP/PCI-glass 898 recycled PET 894 recycled SMC 890,892,893 regrind RIM 895 S2-glass 383 Sglass 503 epoxies 252 single-yam 251 skinned molded foam 258 Skybond 86 SMC 386,386 Spectra 225,503 steel 270 SVM 208 Technora 209 Teijinconex 209 Tekmilon 225 thermoplastic resins 222 titanium 270 Twaron 208 Vectran 235 vinylester 522 VMN-4 305,306 yam distribution 254 ZMC 386,386 stress PP and granulated SMC 897 recycled NBC 897 recycled PP 895,896 Tension and compression 786-9 failure in 787 head, fasteners 625 tension ratio 1049 TEOS 346 testing 781,782, 788 rings 782,788 Tensioning 469 TEOS 340
carbon-carbon composites 349 compression 346 flexure 346 comer 346 impact 346 impregnation 341,3434,344 interlaminar shear 346 interlaminar tension 346 process 341 shear 346 tension 346 thermal expansion 346 Terephthalic acid 35 Terephthalics 36,39 processing time 36 reactivity 36 solubility 36 Tertiary amines 104 Test direction, yam distribution 154 methods 721 adhesives 630 joints 626 specimen dimensions 786-7 loading sections 786 transition sections 786 Testing barcol hardness 38 bars 784 compression 38,782 distortion temperature 38 eddy current 849-50 flat specimens 786-7 flexural properties 38 half-discs 782 history 779 hydraulic system 782 impact strength 38 in-plane shear methods 783 incoming materials 20 microwave 853 non-symmetric bending 783 off-axis tension 781 pressure compliant ring 782 rail shear 783 requirements 779-80 ring specimens 782,784 rods 784 selecting techniques 779 special tests 792 specimens 780,786 splitrings 784 structural 792 tensile properties 38 tension 782 torsion of square plate 783 tubes 784 ultrasonic through-transmission
Index 1113 8404,841 Tetraethylenepentamine (TEPA), viscosity 54 Tetraethylorthosilicate, see TEOS Tetramethylbisphenol F homopolymer property 102 structure precursor 102 supplier 102 tradename 102 Tex, definition 136,241 Textile fibers 307 prefonning definition 397 linear 398 planar 398 three-dimensional 398 preforms 401-2 role of 398-401 selecting 401 T, 120 NR-150 82 PMR-15 82 polyarylene ether 121 polyester 121 polyimides 121 polysulfones 121 resins 440 Thermal accelerators 509 coefficient BMC 386,386 Kevlar 383 LPMC 386,386 ZMC 386,386 coefficients,orthotropic layers 742 conductivity ABS 259 aliphatic system 70 aluminum 558 aromatic system 71 carbon fibers 169 carbon-carbon composites 333,349 carbon/epoxy prepreg 558 cellulois acetate 259 ceramics 307 electroformed nickel 558 epoxies 259 glass/epoxy 558 high carbon cast steel 558 Invar 558 Kevlar 206 mahogany 558 monolithic graphite 558 Nomex 206 particulate fillers 242 phenolics 259
polypropylene 259 polyurethane 259 polyvinyl chloride 259 skinned molded foams 259 steel 558 Technora 206 Teijinconex 206 thermoplastics 531-2 Twaron 206 urethane board stock 558 control tape 819 cycling 89,89,816,817 aircraft 818 spacecraft 818 data 818 decomposition, carbonyls 293-4 degradation 31 diffusivity, thermoplastics 534 expansion 3,5,19 adherends 630,661 carbon-carbon composites 333 ceramic composites 312 coefficient continuous fiber 162 staple 162 whiskers 162 composite tools 592 correction, tooling design 590 E-glass 383 graphite 383 molding 593 NR-150 82 PMR-15 82 S2-glass 383 SMC 386,386 steel dies 507 TEOS 346 thermoplastics 127 tooling materials 577 extremes 816-17 gradients, carbon-carbon composites 333 inertia technology 444 mismatch 638,661,662 oxidation 79 carbon-carbon composites 345,347-8 properties aramid fibers 2057,206 BMI 106,209 carbon fibers 1845,185,186 CE 106,109 composites 661 fiberglass 135 rotational molding 551 radiation 816 resistance, aramid fibers 205 stability, composite tools 592
stress 659,660,661-2 laminates 322 Thermally induced stresses 640 resistant gloves 829 Thermoelastic behavior, anisotropic bodies 470-1 Thermoforming 436,437,526, 529-30,529,532,5459 polymer properties 545 Thermoplastics 356 aerospace 115 aging 127 amorphous, HDT 126 automotive market 115 chemical resistance 126 composite density 534 composites extrusion 526 injection molding 526 processing 525-52 compression molding 116,544-9 compressive properties 124,125 diaphragm forming 116 dielectric constant 107 dissipation factor 107 extrusion 534-8 fatigue resistance 115 fiber content 117 folding 116 fracture toughness 124 heat capacity 5 3 2 4 high price of 1024 impact energy 125 injection molding 116 joining 127 low toxicity 115 manufacturing 116-17 matrices, aramid fibers 221 melt, high viscosity 116 molecular weight 116 polymers adducts 526 fillers for 527 reinforcements 526 prepregs 357 processing 115 properties 115 pultrusion 116 reinforcing 528 remelting 115 resins 31-2 fracture toughness 122 health effects 834 injection 116 tensile modulus 222 tensile strength 122 solvent resistance 115,127 structure 118
1114 Index suppliers 119 thermal conductivity 531-2 diffusivity 534 expansion 127 thermofonnbg 544-9 tougheners 106 trade names 124,125 transverse modulus 123 transverse strength 123 unidirectional compression strength 125 tradenames 125 winding 458 Thermosets amorphous 118 compressive strength 124 joining 127 polymers 34 Thermosetting 37,356 binders 458 dielectric constant 107 dissipation factor 107 reaction, definition 49 resins 518 Thermostamping 22 Thick composites, consolidation 585 Thick molding compound, see TMC Thick-walled structures 456,468-71 loaded in torsion 473-4 monolithicity 456 residual stresses 469 torsion strength 473 Thickeners 383 Thickness adherends 627-9,628,628,658 bond 631 layer 628,656 toadherend 641 drawings 731 fiberglass 141-4,153 Hercules IM7 fiber 478 honeycomb 263 preform 438-9,438 sandwich 284 section 512 variation 372-3 Thin composites, consolidation 585 Thin-walled beams 747-9,748 shear stress resultants 748, 748 uses 747 shells 456 structures 466-8 Thinning, organic solvents 65 Thixotropic liquids, adhesives 275-6 Thomel 1054 properties of 170
Three zone model, pultrusion 505, 505 Three-dimensional fiberglass 146 lamina properties 761 reinforcement 456,801 Threshold limit values, see TLV Through-the-facesheet evacuation, aircraft 865 Time weighted average, see TWA Time window, consolidation 588 Titanates, coupling agents 251 Titanium alloy, linear thermal expansion 705
carbide 319 density 170 diboride 319 elastic modulus 170 fracture elongation 170 specific modulus 1049 specific strength 1049 tensile strength 170 tension-tension ratio 1049 TLV 825,826 Tm polyamide-imides 121 polyarylene ether 121 polyester 121 polysulfones 121 Th4C 380 Tolerance stackup 561 to flaws, ceramics 307 Tonox 60-40 (Uniroyal Inc) 56 Tool geometry 439 life 599 factors 5967 pins 574 separating from master 570 stee1,CTE 589 surface area 480 strength 480 usage 559 wear 599 Tooling 14,361-6,392-5,429-30, 447,47940,480,556-75 alternative 449-50,449,450 balls 565 bushings 562 castaluminum 442 correction method 590-1 CTE! 5567,5574 design 590 coordinating partial plies 590 thermal expansion correction 590
use of caul plates 590 development, space 1015-16 elastomer 14 elastomeric 5934,593 electroformed nickel 591-2,592 employee induced damage 563 epoxy laminates 562 fiber placement 479-80 fiber separation 562 fixed-volume method 593,593 graphite-epoxy 592-3,593 hole fittings 562 leakage 562 master models 563-6 materials 589 bleeder 577 consolidation 577 hardness 577 metals 577 properties of 558 stiffness 577 strength 577 surface roughness 577 thermal behavior 589 thermal expansion 577 moisture absorption 561 and part shape 589-94 prepreg 562 resin system 562 rubber 366 sealers 561 selecting materials 557 sheet wax 574 solidrubber 670 spray metal 442 urethane based 567 variable volume method 593,594 wet lay-up repairs 867 Tools composite 361 curecycletimes 559 depth of cut 597 design 556-63 durability 558 fabrication, cost 558,559-60 life expectancy 558,559,561 machinability 558 materials available 560 materials selection 559 methods of fabrication 560 mounted 480 multi-cavity 500 resistance to failure 561-3 selection 598 surface finish 597 usage rates 558 wear ratio 597 Torayca 1054 fiber types 1054
Index 1115 properties of 170 Torque levels 625 bolt tensions 626 Torsion 784,790 square plate, testing 783 strength, thick-walled structures 473 Total hip arthroplasty 960-1 femoral components 960-1 FRI' 959 metallic devices 960 Total overlap length 658 Tougheners reactive rubber 106 thermoplastic 106 Toughening 104-5 interphase 802,802 Toughness 801 adhesives 2765,274 aramid fibers 221 interlaminar 124,124 measuring 274,274 resins 440 Tows 164 bridging of 487 buckled 480 carbon fiber 292,293 cutting 483 definition 476 dropping and adding 4834 end locations 481 folded 487 selecting 478 thickness 478 twisted 487 wandering 487 width histogram 479,479 of material 479 Toxic 8 2 2 4 definition 822 Toxicity acute and chronic 823-4 fluorine 362 particulate fillers 252 reproductive 824 solvents 361,362 Trade names continuous fiber 162 fibers 158 PAN-based tow 198 pitch-based tow 199 polyamide-imides 121 polyarylene ether 121 polyester 121 polyimides 121 polysulfones 121 staple 262 thermoplastics 124,225
whiskers 158,162 Transfer molding 22 Transformation relations 707 Transition, areas of 731-2,732 Transportation 905-15 future directions 915 land 90515 pultrusion 520-1 Transverse compression, Kevlar 212 CTE 699 extension, Ih46/epoxy 701 modulus 12 Ih46/epoxy 699 M401/F854 699 resins 1056 thermoplastics 123 shear adherends 630 failure, sandwich 282 strain, IM6/epoxy 701 strength moisture 811 thermoplastics 223 tensile strength, resins 1056 tension Kevlar 212 test 688 Treatments, Vectran 236 Triaxial, plain-weave 399 Triethylenetriamine (TETA), viscosity 54 Triglycidyl p-aminophenol 53 Trimellitic anhydride (TMA), melting point 58 Trimming 446 Kevlar 600 2,4,6-Tris( dimethylaminomethy1)phenyl 60 Trucks 912 Truss structure, spacecraft 968-70 TRW-R-8XX 94 chemistryof 81 Tubes, testing 784 Tubular structures 425-32 pultrusion 496-7 Tunnels 950 Turning 597,598,601 advantages 597 withAWJ 603-4 disadvantages 597 Kevlar 600 machining 602 repeatability 603 surface finish 603-4 TWA 826 Twaron availability 219-20 decomposition temperature 206
density 206 elongation at break 208 filament diameter 206 filament shape 206 forms of 216 initial tensile modulus 208 linear thermal expansion coefficient 206 long-term use temperature 206 melting temperature 206 moisture content 206 properties 205 sources of information 221 specific gravity 208 specific heat 206 tensile strength 208 thermal conductivity 206 Twill, weave 145,145,405,406 Twin-screw extruders 535-6 advantages 539 classification 537 disadvantages 539 Twist drill 600 Kevlar 215-16 Two-dimensional flow, compression molding 584 lamina properties 760-1 reinforcement 456 Typical compression strength, carbon fiber composites 1055 properties 1053-8 E-glass 1055 selecting 1057-8 Tyranno 316 UHMWPE 959,962 compression molding 545 reinforcement 963 Ultimates, Kevlar 210 Ultrahigh molecular weight polyethylene, see UHMWPE Ultrasonic correlation 8434,844 cutting 601 machining(USM) 6058 advantages 608 applications 607-8 brittle materials 605 ceramic matrix composites 605 disadvantages 608 rates 608 polar backscatter 843 pulse-echo testing 842-3,842 resonance 843 signal 840 tonoiseratio 834
1116 Index through-transmission testing 840-2,841 automation 840 C - S C 841,841 ~~ vibrations 607 Ultraviolet Coronagraph Spectrometer, see UVCS Ultraviolet resistance, PET 224 UMC 381 Uniaxial loading, allowable 802 Unidirectional composites 107-8 compression testing 787 fabrics, weave 145,145 laminates 4,8 molding compound, see UMC ply 741 tape, flexural strength 401 Uniformity 301,301,302 Unit costs, polymer processes 532 Unsaturated polyesters, resins 382 Unsaturation 39 Unsupported films, adhesives 276 single lap, adhesive joints 628 Untapered, joints 624 Uranium enrichment centrifuges 953-4 Urea-formaldehyde, adhesives 255 Urethane board stock ceofficient of thermal expansion 558
density 558 thermal conductivity 558 US Air Force Materials Laboratory 22-3 US Federal Aviation Regulations (FAR) 1022 Use temperature, PET 225
uv
absorption, aramid fibers 221 radiation, carbon black 245 resistance, PET 230,233 stability PBOfiber 238 PET 233 UVCS 1017-18,1018 V-cap 93 chemistryof 81 processing conditions 84-5 Vacuum assisted resin injection, see VARI assisted resin transfer molding, see VARTM bag 14,14 autoclave tooling 589 molding 361,577 nylon 577
pressure 918 processing 919-20 silicone rubber 577 bagging 569,593,872-3 applications, aerospace 375 with autoclave 593 bridging 369-71,369 eliminating 369,370,371 edge bleeder 368 expendable 367-71 bags 371 precautions 363 reusable 371-2 thermoset composites 371 vertical bleeder 368 drying 818 infiltration 299-300 pressure cures 870 repairs 869 Valves 954 Vapor barriers 817 processing 316-17,316 VAN 910 VARTM 919 advantages 919 Vectran 234-7,234,235 aerospace applications 236 applications 236-7 availability 236 cost of 234-5 dielectric constant 236 elongation at break 235 forms 236 melt spun 235 melting temperature 235 properties 235-6 protective garments 236 resin 235 safety materials 236 sources of information 236 specific gravity 235 tensile modulus 235 tensile strength 235 treatments 236 Veils 436,438 surface ply 503 surfacing, application 436 Velocity, meteoroids 815 Velsicol 58 Venetian blinding 482 Vent design 442-3 Ventilation 830 volatiles 683 Venting 450-1 ports 443 Verification, design 733-4, 733 Vertical
pultrusion 495-6 advantages 495 tape wrapper 427,428 Viiyl cyclohexene dioxide, viscosity 63 Vinyl esters applications 41-2 corrosion resistance 41,42,42 density 505 elongation at break 505 flexural modulus 505 flexural strength 505 general purpose 40 heat distortion 505 heat resistant 40-1 marine applications 916 properties 41-2 resins 28,382,504 structure 40 tensile modulus 505 tensile strength 505 Viscoelasticity 122 Viscosity 37 aliphatic system 70 anhydride-cured system 72 aromatic system 71 coupling agents 250 definition 527 elongational 529-30 low density polyethylene 528 polypropylene 530,531 predicting 528-9 resins 440,578 RTM 440 shear rate-dependent 528 and temperature 173-4,174 vs. time, epoxy resins 509,510 Visual inspection 839-40 VLS 160 timeneeded 160 vMN-4 bending strength 305 carbon content 306 coating thickness 306 compressive strength 305 density 305,306 elastic modulus 305 elongation at break 306 filament diameter 306 specific modulus 305 strand properties 306 tensile modulus 306 tensile strength 305,306 Voids 430,548,588,630-1 capture 450-1 content checks 430 marine laminates 920 exposed surface 432
Index 1117 stability map 588, 588 Volatiles content 356 elimination 78,79 ventilation 683 Volkersen model 662-3 solution 635 Volume expansion, water absorption 811 fraction fiber reinforcement 435 IM6/epoxy 699 M401/F854 699 resistivity fiberglass 235 Skybond 95 Volumetric shrinkage aliphatic system 70 aromatic system 72 Von Mises criterion 708 Vortex, tennis rackets 1046 Voyager aircraft 32, 32 Walleffects 153 Warp 541 knitting 408,409 plied-yam 252 single-yam 252 yam 140 fiberglass 141-4 Warpage 374,467,557 Warping GFRP 1013,1013 polyoxymethylene 541 Wash resistance, fiber reinforcement 435 Wastage cut-off 502 pultrusion 489 Water absorption 110,811 aliphatic system 70 aramid fibers 215 aromatic system 72 rate of relaxation 811 recycled NBC 897 recycled PET 894 recycled SMC 890 swelling stresses 811 volume expansion 811 evacuation, aircraft 863 immersion 110 ingestion, aircraft 859 intrusion, NDE 854 jet 602 resistance 3 soluble sand mandrels 465-6 Weak bonds 671-5’671
Weather resistant, resins 813 Weathering 813 climate 813 graphite/epoxy 813 graphite/polysulfone 813 Kevlar/epoxy 813 Weave aramid fibers 222 basket 140,145 crowfoot satin 145,145 eight-hamess satin 145,145 fiberglass 2414 non-woven 145,146 patterns 140,145,150 and mechanical properties 2 52 plain 140,145 biaxial 407 plied-yam 252 shearing 439 single-yam 252 twill 145,145 unidirectional fabrics 145,145 Weavers 407,407 Weaves hybridized 128 reinforcing 118 Weaving 18,402,402,4044,405, 406,407,801 biaxial 406 plain weave 405 satinweave 405 twill weave 405 definition 404 design methodology 407-8 multiwarp 406 loom 405 three-dimensional 405,405 triaxial 405, 405,406 two-hamess loom 405 Web, yams 407,407 Weft, knitting 408,409 Weibull distribution 189 Weight aramid fibers 223 fiberglass 142-4 fraction 1027-8 loss AFR700B/S2 laminates 94 PBOfiber 238 Skybond 86 savings 1028,1028 Weld lines 542,544 strength 544 Wet lay-up 355,571-3 composite tools 566 repairs aircraft 865-7
core plug 866 damage assessment 866 elevated temperatures 868 equipment 867 face sheet 866-7 heat lamps 865,866 heating blankets 865 hot air blowers 865 materials 867 precured patching 867-8 procedures 866-7 repair environment 867 restoration of coatings 867 tooling 867 tworesins 571 Wet mat molding 43 Wetout 443 Wettability, fiber reinforcement 435 Wetting agents, additives 516 Wetting speed 495 Whipple-type shield 814-15,814 Whiskers 160-1,308,308,308,314 alumina 161 coating 166 commercially available 308 composition 258 continuous 156,160 CTE 308 density 262,308 description 262 diameter 262,308 discontinuous 156,160-1 elastic modulus 162 length 308 manufacturers 258,262 manufacturing processes 258 for reinforcing 166 Sic 308 elastic modulus 327 fracture strength 325,326, 326,327 fracture toughness 325,326, 326,327 silicon nitride 161 specific strength 262 thermal expansion coefficient 262 trade names 258,262 vapor-liquid-solid process 160-1, 326 Young’s modulus 308 Whole rings, bending 785 Width of material 479 Wind eye 461 tunnel blades 975-6,976 turbine blades 954 Winders lathe type 461 race-track 462
1118 Index stationary whirling arm type 461 Winding angle 496-7 with braiding 473 chord 462 direction 496-7 dry 458 extra pressure 470,470 geometry 459 helicalangle 460 history of 458 inaccuracies 463 isotropic metallic tape 468 layer-by-layer curing 470 materials for 458-9 non-geodesic 457 pattern 459 closure 459 control 4 6 3 4 pitch of 496-7 planar-polar 462 preform 458 process 456,457 programmed 469-70 pull 489,490 and pultrusion 496-7 semifabricated composite 468 speeds 463 stages 469 technology 458-66 theory of 468-9 thermoplastics 458 of a torus 462 wet 458 Windmills 521 Wood 256,838,958 durability 256 end-grain balsa 256 low Cost of 256 rehabilitation 988 Work of fracture alumina composites 325
Silar 325 Tateho 325 Work-around technique 1006,1012 Workmanship 715 Woven fabric 359 laminates 818 flexural strength 401 reinforcements 91&17 Wrapping tapes 429 Wrinkles 487 X-ray backscattering 8454,846,547 imaging 844-9 cost advantages 845 foreign materials 844 honeycomb defects 845 matrixcracks 844 moisture contamination 858 non-film 845 porosity 844 XMC 381 Yams classification 402 content, distribution 154 crimp 416 cross sections 411 designation, fiberglass 23940 distribution 154 filament 164 fill 407,407 filling 140 geometry 4034,403 identification 138-9 inclination angle 407 jamming 412 longitudinal 415 orientation angle 412 Stwist 138
slippage 399 surface weave 407,407 textured 138 twisted, geometry 404 warp 140 web 407,407 Z twist 138 Yield fiberglass 139-40 stress at break, PP/PCI-glass 898 Young’s modulus alumina composites 325 aluminasilicate 309 carbon fibers 295 CE 108 ceramic composites 312 ceramics 307 E-glass 24 and microtexture 189 resins 440 Silar 325 Tateho 325 whiskers 308 Z molding compound, see ZhtC Z twist, yam 138 Zero-shear viscosity, rotational molding 552 Zirconates, coupling agents 251 ZMC 380 flexural modulus 386,386 flexural strength 386,386 formulation 384 injection machine 380 IZOD 386,386 low viscosity 380 specific gravity expansion 386, 386 tensile modulus 386,386 tensile strength 386,386 thermal coefficient 386,386 Zonyl 251
