U.S. DEPARTMENT OF COMMERCE National Technical Information Servi;e
AD-A033 216
ENGINEERING DESIGN HANDBOOK HELICOPTER ENGINEERING.
PART TWO
DETAIL DESIGN
ARMY MATERIEL COMMAND,
ALEXANDRIA,
JANUARY, 1976
Reproduced From Best Available Copy
VIRGINIA
351058 AMCPAMPIT
AMCP 706-202'-
ENGINEERING DESIGN
HANDBOOK HELICOPTER ENGINEERIN\
PART TWO DETAI L DESIGN
IIEADEQMRE-USP 1 AINYV MATERIEL COMMAND NATIONAL TECHNICAL INFORMATION SERVICE U.S. DEPARTMENT OF COMMIE," iMNWacw, VA. 2231W
JANUAR IC76
AMCP 706-202
AN1( Parrnpbkei
No. 714S-01
ENGINEERING I)ESIGN HANI)BOOK 20Jnay17 HELCOPERENGINEERING, PART TWO DETAIL. DESIGN
TABLE (N-' CONTENTS Paragraph
EE LISTOF ILL.USTRATIONS................ ......................... I 1ST OF TABLES ... .... .... ................................... FOREW ORD .. .. .. . . .. . . .. . . ....... . . .. . . . . . .Xxxviii PREFACE ... . . . . . . . . . . . . . . .
Xxviii
Xxxi, XXX%
CHAPTEKR I INTROI)1UCTION (HAPTERI
2-I
)
2/
INTRODULCTION ....... M.Al.'E. A.I..........S.
. . .-
2-2 ~METALS .................. ............. ..... .... ....... ...... 2-1 2-2.1 FERROUS METALS.............................................. 2-1 2-2. 1.1 General .................................. ......... ............. 2-1 2-2.1.2 Ca!'bOn Steels .............. -............. ...... ................ 2-1 2-2.1.3 Allov Steels.........................-2 2-2.1.4 StainessSteels....................... ............................ 2-2 2-2.1.5 Precipitation Hardening Sicels ...................................... 2-2 2-2.1.6 Maraging Steels .................................................. 2-3 NONFERROUS METALS ............................. .......... 2-4 2-2.2 2-2.2. 1 General.............................. ........................... 2-4 2-4 2-2.2.2 Aluminum Alloys........... I................I.................. 2-5 Magnesium Alloys....I............................................ 2-2.2.3 2-2.2.4 Titanium Alloys......................................... ......... 2-6 2-2.2.5 Copp.,r anid Copper Alloys ......................................... 2-6 2-7 2-2.3 ELECTROLYTIC ACTION OF DISSIMILAR MFTAI.S................ 2-3 NOMETALLIC MATERIALS........................................2-4-7 2-7 ............................ GENERAL .......................... 2-3.1 22-3.2 THERMOPI ASTIC MATERIALS ................................ 2-3.3 THERMOSETTING MATERIALS......... ................ ........ 2-9 ...... 2-10 2-3.4 ELASTOMERIC MATERIALS ............................. .. ........... 2-10 2-3.5 WINDOW MATERIALS ............................ COMPOSITE STRUCTURES.........................................21-Il 2-4 .. ........... 2-Il FIBERGLAS LAMINATES .......................... 2-4.1 2-4.1.1 Design Considerations ................ ............................ 2-1l 2-4.1.2 Resin Systems ................................................... 2-12 2-12 2-.Ž;Polyesters ...................................................... 2-4.1.2.2 Epoxies ........................................................ 2-12 2-4.1.2.3 Phrnolics .......................... ............................ 2-12
AMCP 706-202 TABLE OF CONTENTS (Continued) Paragraph
Pagc
2-4.1 3 2-4.1.3.1 2-4.1.3.2 2-4.1.3.3 2-4.1.4 2-4.1.4.1 2-4.1.4.2 2-4.1.4.3 '2-4.1.5 2-4.2 2-4.2.1 2-4.2.2 2-4.2.3 2-4.2.4 2-4.3 2-4.3.I 2-4.3.1.1 2-4 .3.1.2 2-4.3.1.3 2-4 .3.1.4 ,,•
Types of Reinforccment .............. ........... .................... Nonwoven Continuous Filaments .................................... Woven Fabric .................................................... Chopped Fiber ..................................................... Fabrication M ethods ................... .............................. O pen Mold H and Layup .............................................. Sprayup ................ ........................... ...... Matched D ie M olding ................................................ Surface F inishes ....................................................... FABRIC LAM IN ATES .................................................. Reinforcem ent Selection ............................................... R esin Selection ........................................................ Special T ypes ......................................................... Specifications ......................................................... FILAMENT COM POSITION .......................................... Types of Reinforcem ent ............... ............................... E-glass ................................................... S -g lass ...................................................... ........ Boron Filam ents ................ ....................... ........... .............. 11lG ..rap hite .............................................
, C
"" ,..
.
.
......
......................
2-4.3.3
Manufacturing Processes ..................... ....
2-4.3.4
A pplications ...................
2.44 2-4.5 2-4.5.1
...................
........................
2-5 2-5.1 2-5.1.1 2-5.1.2 2-5.1.3 2-5.2 2-6 2-6.1 2-6.2 2-6 .3 2-6 .4 2-7 2-7.1 2-7.2 2-7.3 2-7.4 2-7 5
1t
2-19
2-19o .
,
2-20
HONEYCOMB AND SANDWICH CONSTRUCTION . A R MO R MATERIALS ................................................ Available Materials .......................................... Design ADHESIVES AND SEALANTS ........... ........................ BO N D ING AG EN TS ................... .............................. Structural Adhesives ............... ............................... Nonstructural Adhesives ....................................... Processing O perations ................................................. -Dcsgi, ofonded Structurc......................................... SEAL.ING CO MPO UN DS ............................................. PA IN IS A N D FIN ISH ES ................ .... .......................... PAINTS AND COATINGS (ORGANIC) ............................... SPECIAL. FINISHES ............................................... P LAT IN G ......................................................... ... T A P ES ........................... ................ ................... LUBRICANTS, GRI-ASES. AND HYDRAUI.IC FLUIDS ............... G E N ER AL ......... ............................ ..................... DESIGN OF LUBRICATION SYSTEMS ............................ G RE-A SE S ...................................... ...................... DRY FILM AND PERMANENT LUBRICANTS ..................... H YD R AU LIC FI.U IDS ....... ............... ....................... R EF ER E NC ES ...................................................... ...
""4
2-12 2-13 2-13 2-13 2-13 2-14 2-15 2,15 2-15 2-16 2-17 2-17 2-17 2-17 2-17 2-18 2-18 2-18 2-18 2-18
2-20 2-272-29 2-30 2-30 2-30 2-30 2-32 2-33
•
2-33
2-33 2-34 2-34 2-35 2-36 2-3 7 2-38 2-3K 2-38 2-38 2-38 2-40 2-40
(0i1A'ITR 3 PROPII. ION SIBSYSTEM I)IESIN'N
L IST Of: SY M BO LS .....................................................
3-0
3-1
ii
.. . .. .
.
.
.
..
r-
.
.
.
.
.
.
..
~
.
.
.
.
.
.
"AMCP 706-202 TABLE OF (ON1I EN'IS(( ontinud) Page
Paragraph 3-1 3-2 3-2.1 3-2.1.1 3-2.1.2 3-2.1.3 3-2. 1.4 3-2.2 3-2.3 3-2.4 3-2.4.1 3-2.4.2 3-2.5 3-2.5.1 3-2.5.2 3-2.5.3 3-2.5.3.1 3-2.5.3.2 3-2.5.3.3 3-2.6 3-2.6. 1 3-2.6.2 3-2.6.2.1 3-2.6.2.2 3-3 3-4 3-4.1 3-4.2 3-4.2.1 3-4.2.2 3-4 .2.3 3-4.2.4 .3-4 2.5 3-4 3-4 .22.6 .7
"3-4.2. 3-4 .3 3-5 3-6 3-7 3-8 3-8.1 3-8.2 3-8.2.1 3-8.2.2 3-8.2.3 3-8.2.4 3-8.2.5 3-8.3 . .3. I
3-
INTRO D UCTIO N .....................................................
3-1
ENGINE INSTALLATION .............................................. G EN E R A L ............................................................ Subm erged Installation ................................................ Sem icxposed !nstallation .............................................. Exposed Installation ................................................... D esign C hecklist .......................................... .......... ENGINE M OUNTING ................................... ............ ENGINE VIBRATION ISOLATION ................................... FIF EW A LLS .......................................................... Fire Detectors ......................................................... Fire Extinguishing ..................................................... ENGINE AIR INDUCTION SUBSYSTEM ............................ Air Induction Subsystem Design ....................................... Inlet Protection ....................................................6.. A nti-icing ............................................................. Electrical A nti-icing .................................................. Bleed Air Anti-icing ............................................... Anti-icing D em onstration ............................................ EXHAUST SUBSYSTEM .............................................. ....... ............... Exhaust Ejectors . .......................... Inftaid (IR) Radidtiui Suppr, .sior ................................... 1R Suppression Requirements . ...................................... Exhaust Suppressor ................................................. PROPULSION CONTROLS ............................................. .................. ... ....... FU EL SU BSYSTEM ................... ................................... ... GENERAl . ............. FLUEL SUBSYSTEM COMPONENTS ................................ Fuel Tanks ................................................. ........ . Fuel Tank Venis ....................... .............................. . F uel G aging ........................................................... Refueling and D efueling ............................................... F uel D um ping ........................................................ ................... Engine Feed System .............................. Fue l l)r iti ....................... ..... .... .........................
3-i 3-1 3-1 3-I 3-3 3-4 3-4 3-5 3-5 3-5 3-6 3-6 3-6 3-6 3-7 3-7 3-7 3-7 3-7 3-8
....................... Controls and Instrumenlation ................ T EST IN G ............................................................. LUBRICATION SUBSYSTEM ......................................... COMPARTMENT COOLING .............. ACCESSORIES AND ACCESSORY DRIVES ............................ ................................ AUXILIARY POWER UNITS (AIU', G EN ERA L ......................................................... APU INSTALLATION DETAILS ..................................... Method of Mounting .................................................. Inlet D ucting ........................ ............................. ... Exhaust Ducting .................................................. ................................ APU Bleed Air Ducting ........ Cooling ..................... . .. ....................... . ............. APU SU BSYSTEM S ....... ......................................... .............................. C ontrols . ...................
3.13 3-13 3-14 3-14 3-15 3-15 3-15 3-15 3-15 3-16 3-17 3-18 3-18 3-18 3-18
..............
Electrical
3-8
3-1 3-9 3-9 3-9 3-9 3-10 3-10 3-11 3-11 3-11 3-13 3-13 3-13
Iiii
AMC? 7W0620
___
T'ABI.ELF
(ONr'' "I %I
i( onalinucd)
Paragraph
3-8.31. 3-8v .1.2 3-8.3.1 3 3-8.3.1.4 3-8.3. 1.5 3-8.3.2 3-8.3 2.1 3-8.3.2.2 3-8.3.. 3-8.3.4
3-8.3.5 3-8.4 3-8.5
____
Pdgv
Sequencing Controls ................................................. Protective Controls ....... ............................ .............. O utput C ontrols ..................................................... Electrical Control Location ....................................... ... Electrical Power Requirements ...................................... Fuel System Controls .................................................. Rated Speed G overning .............................................. Filtering Requirements ............ ................................ APU Lubrication Subsystem .......................................... A PU Reduction Drive ..... ............................... ...........
3.18 3-18 3-18 3-19 3-19 3-19 3-19 3-19 3-20 3-20
AP Starting.................
3-20
.................................
R ELIA BILIT Y ........................................................ SAFETY PROVISIO NS ..................................... .......... R EFER EN C ES ..........................................................
3-20 3-21 3-22
CHAPTER 4
TRANSMISSION AND DRIVE SUBSYSTEM DESIGN 4-0 4-I 4-1.1 4-1.2
4-1.2.1 4-1.2.1.1 4-1.2.1.1.1 4-1.2.1.1.2 4.1.2.1.1.3 4-1.2A. 1.4 4-1.2.1.2 4-0.21.3
"4-1.2.1.4 4-1.2.1,4.1 4-1.2.1.4.2 4-1.2.I 4.3 4-1.2.1.4.4 4-1.2.1.4.5 4-1.2.2 4-1.2.2.1 4-1.2.2.2 • : 4-1.2.2.3 4-1.3 4-1.3.1 4-1.3.2 4-1.3.3 4-1.4 4-1.4.1 4-1.4.2
"4-1.4.2.1 4-1.4.2.2 4-1.4.2.3 iv
liSTOF SYMBOLS .............................................. INTRO D U CTIO N ....................................................... G EN ER Al. ........................................................ ... REQUIREM ENTS ................................................... . G eneral Requirements .... ......................................... Pe-fortrance ......................................................... Subsystcni Weight ........................ ..................... Transm ission Efficiencr y ................................. ........... Si Ce Levels. ..................... ........ .......................... N oiseLevels ....... . .............................................. R eliability ..... ................................................. Maintainabifity.................................................. Survivability ............................................... R edundancy ........................................................ Dcsigu Configuration ............................................... Self-sealing Sum ps .................................................. Emergency Lubrication ......................... ............ ....... A rm or .............................................................. Drive System Configurations ......................................... Single M ain Rotor Drive System ..................................... M ultilifting-rotor Drive System s ...................................... Compound Helicopter Drive Systems ................................. TRANSMISSION DESIGN AND RATING CHARACTERISTICS ..... Power/Life Interaction ................................................ Transmission Overhaul Life Rating ..................................... Transmission Standards and Ratings ................................... QUALIFICATION REQUIREMENTS ................................ Component and Environment............... ........................ Dcvelopment Testing ...................... .......................... Static C asting Tests ................................................... Deflection Tests ...................................................... C ontact 1 :sts ........................................................
4-1 4 3 4-3 4.3 4-3 4-4 4-4 4-4 4-I1 4-11 4-12 4-16 4-7 4-I1 4-18 4-22 4-22 4-23 4-23 4-23 4-23 4-25 4-25 4-25 4-27 4-28 4.29 4-2'9 4-29 4-29 4-29 4-30
( .•
4 :.
AMCP 706-202 lARIl.01- (()%'I IKNIS (Oininuvdli
Paraigraph 4-1.4.2.4
4-1.4.2.5 4-1.4.2.6 4-1.4.2.7 4-1.4.3
.L4-1.4.4
4-2 4.2.1. 4-2.1.1 4-2.1 .2 -4-2.1.2.1 4-2.1.2.2 .2.3 4-2.2 4-2.2..1 4-2.2.1.1
It4-2.1
4-2.2. 1.2. 1 A
-Bending Fatigue Strength ............................
...........
ailuir .................................................................
4-30 4-30 4-30 4-31 4-31 4-32 4-32 4-32 4-32 4.34 4-34 4-34 4-34 43 4-34 4-35 4-36
Cattie Failure.................................................
4-44
4-2.2.1.2.3.2
Classic or Pitch Line Fatigue ....................................
445
4-2.2.1.2.3.3
Wear Initiated Failure
446
4-2.2.2.2.3 4-2.2.2.3 A-2.2.3 4-2.2.3.1 4-223.2 4-2.2.3.3 4-2.2.3.4 4-2.2.3.5 -4A-2.2.4
4-2.2.4.1 4-2.2.4.2. 4-2,2.4.3 4-2.2.5 4-2.2.5.1 4-2.2,5.2
e:4-2.3 4-2.3. 1
.........
I...........I.............
r~
44.
4-2.2.1.2.3.
''4-2.2.2.2.2
)
Assembly and Disassembly ....................................... Lubrication System Debugging.................................... Incremental Loading and Efficiency Tests ........................... Thermal Mapping Tests.......................................... Overpower Testing............................................... Other Life and Reliabil~ty Substantiation Testing .................. .... TRANSMISSIONS ................................................ FAILURE MODES .............................................. Primary Failure Moocs ........................................... Secondary Failure Modes ......................................... Overload Failures ............................................... Debi is-caused Failure ........................................... Environmentally Induced Failures ................................. DYNAMIC 1C.CNPONENTS .................................... Gears Limi.....ations............................................ Gear Lnialyios ................................................ aiukgl;I
4-2.2.1.3 4-2.21.2 4-2.2.2. 1. 4-2.2.2..1 4-2.2.2.1.2 4222.3 4-2.2.2.1.4 '4-2.2.2.2 4-2.2.2.2.1
Y
Pp
Gear Drawing and Specification................................... 4-4 Bearings ....................................................... 4-48 Lubrication Deighnq.......s.....................................4-4O AMpluctiong Driesign ......................................... .. 4-48 Mounrctiong Practices .......................................... 4-48 Internal Characteristics ......................................... 45 Skidding Control .............................................. 4-54 Life Anm lysis ................................................. 45 Assumptions and Limitations .................................... 4-55 'Modification Faiflot Approach to Lifc Prediction..............I...... 4-55 Complete Elastic and Dynamic Solutions........................... 4-56 Drawing Controls........................................ ...... 4-56 Splines.........................................................44-7 Face Splines ......................................... ......... 4-386L Concentric or Iongitudinal Splineq ................................. .5 :Propcrtics ofSplineb ............................ ................ 4-58 Spline Strength A nalysis..................... .................... 4-59 Drawing Design and Control ..................................... 4-60) 4-60 -Overrunning Clutches ............................................ Sprag Clutches .................................................. 4-61 Ramp and Roller Clutches........................................ 4-62 Self-encvgizing Spring Clutches.................................... 4-62 Rotor Brakes........................ ........................... 4-62 R,ýquirernents and Limitations.................. .................. 4-62 Design arid Analysis ........................................... 4-63 STATIC COMPONENTS.................................... ..... 4-64 Casez end Housings.............................................. 4-64 v
.
r.
..
.
..
. .
. .
0 AK 70_2o2 TABLE 0 CONTENTS Iedlnutd) P lParagraph
Pug-
4-2.3.1.1
4-4.2 4-4.2. I 4-4.2.2 4-4.3 4-5 4-5.1 4-5.2 4-5.3
4-64 4-66 4-67 4-W 4-0. 4- 3 4-72 4-72 4-72 4-73 4-73 4-74 4-76 4-76 4-76 4-80 4-81 4-81 4-82 4-82 4-83
J 3.rt..C. a.^CO*
44 13.3
--
.................................... .............
M ate:ials and Pro .............................................. Quills ...................................... .......................... SPECIAL CONSIDERATIONS ........................................ Vibr'ation Control ..................................................... D iagnostics ........................................................... DRIVE SHAFTING AND INTERCO iNECTION SYSTEMS ............ GENERAL. REQUIREMENTS ........................................ Engine-to-Transmission ............................................... Interconnect Shafting .................................................. Tail Rotor or Propdlcr Shafting ....................................... Subcritical Shafting ........................................ .......... Supercritical Shafting ................... .............................. COMPDONEi-NT DESIGN ............................................. C ouplings ............................................................. Bearings ................................................... Shafting .............................................................. LUBRICATION SYSTEMS ....................................... OIL MANAGEMENT ............................................ Function .............................................................. Comporent and Arrangement .........................................
Desi.5n ai'd Analys-u
4-2.3.1.2 4-2.3.2 4-2.4 4-2.4.1 4-2.4.2 4-3 4-3.1 A-3. I. I 4-3.1.2 4-3.1.3 4-3.1.4 4.3.1.5 4-3.2 4-3.2.1 4-3.2.2 4-3.2.3 4-4 4-41 4-4.1.1. 4-4.1.2 .
..1# 0
.
.
............
......
COOLING REQUIREMENTS ......................................... Heat Exchanger Sizing ................ ................................ Cooling Fan Sizing ........................................... EMERGENCY LUBRICATION.................................. ACCESSORIES......................... ............................ PAD LOCATION AND DESIGN CRITERIA................ ......... ACCESSORY DRIVE DESIGN REQUIREMENTS .................... REQUIREMENTS SPECIAL F.........................................
4s
4-86 4-36 4-87 4-87 4-88 4-88 4-89 4-89
QC
(
,2,-.
K
CHAPTER 5
ROTOR AND PROPELLER SUBSYSTEM DESIGN
vi
5-0 5-1
LIST OF SYM BOLS .................................................... IN TRO D UCTIO N .............................................. ........
5-1 5-2
5-2 5-2.1 5-2.1.1 5-2.1.2 5-7-!.3 5-2.1.4 5-2.1.5 5-2.1.6 5-2.1.2 5-2.1.8 5-2.2 5-2.3 5-2.4 5-3 5-3.1
DESIG N PARAM ETERS ................................................ HOVER Disk Loading and Induced Power ...................................... Blade Loading ....................................................... Blade Tip M ach N um ber ....................... ..................... N um ber of Blades ......................... ........................... Tw ist .............................................. .................. A irfoil Sections ......................................... .............. Hovering Th vust Capability ............................................ ........................................... Guideines .......p HIGHu SPEED LEVEL F. IGHT ...................................... HIGH-SPEED MANEUVERING FLIG:T ........................... IN E R T IA .................................. ........ .................. ROTOR SYSTEM KINEMATICS ....................................... G EN ER A L ..................................................... ... ...
5-3 5-3 5-3 5-4 5-5 5-5 5-5 5-5 5-5 5-5 5-6 5-6 5-7 5-7 5-7
(
-W
W4
706-j22
_________AlMiýP
SA OLL OF (CONTENTh I(vatinuLedI PararaphPatic
5-3.
11 LICOPTER CONTROLa.......................
5.9
5-3.j
ARTICUL,%TID R0104.........................
59
5-3.4
GtNMBALED(TEI-TERING)HO1TOR
5-10
5-3.5
HIINGELESS ROY OR .........................
5-3.5.1 5-3.5.2 5-3.6 5-3.6.1 *
-5-3.6.2
5-3.6.3 5-4 5-4.1 5-4.1.1 54.1.2 5-4,1.2.1 5-41.2.2) 5-4.1.2.3
-.
Fr~tigue Tests ....................................................
5-4.3
GROUND RES.NONANCE
f
>
AS
rrQ
wlS.................................
To-blded otorWithHinged Blades ............................... Two-bladrd Riatocs WVitbout Hinges .................................. M ultibiaded Rotors ........ .... .. ... ......... FLUTT FR ASSESSMENT .......................................... Current Criteria................................................... Design Considerations ............. ..................... .......... Helicopter.............. ........ ...................... ......... 5-4.4.2.1.1 FaAed SystEm ................................................... 5-4.4.2.1.2 Ro'ating System ... ............................................ 5-4.4.2.2 Compound ...................................................... 5 -4.4.2.2. i Fixed System ................................................... 5-4.4.2..72 Rotating System ................................................ 5-4.5 ACOUSTIC LOADING ............................................ 5-4.6 GUST LOADINGS................................................ 5 4.6.1 Discussion of the Gust Prohi -m...................................... 5-4.6.2; Guist Design Considerations......... ............................... 15-4.7 TORSIONAL STABILITY ............................. ............ 5-4.7. 1 Discussion of Problem . ...... .. . . . .. ... ........ 5-4.7.2 De!sign Considerations ....... .. . . . ... .. ........ 5-5 BLADE RETENTION .......................... 5-51I RETENTION SYSTLM DESIGN CONSIDElRATIONS................. 55 .1Articulated Rotors ... . . .. . .. . . . . . . .. .. . .. . . . Articluated Rotor Considerations ............... -51.1Typical 5-5.1.1.2 Reversed Hinge Articulation ...................... 5-5.1.2 Gimbaled and Teetering Rotors ... ................. 5-5.1.2.1 Gimbal-mounted Hubs .......... .............. 5-5.1.2.2 Teetering Hubs . . . . . . . .. . . . . . . . . . . . . . . . . .. . 5-5.1.3 Rigid Rotor . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 5-5.2 COMPONENT DESIGN CONSIDERATIONS ............. 5-5.2.1 Rolling Element Bearings ..... .................. 5-4..1
5.4.3.2 5-4.3.3 5-44 5-4.4.1 5-4.!.2 5-4.4.2.1
5-11
XH-51 Rotor System .......................... 5-12 OH-6A Rotor .. .... ... .. . . .. . . ... .... ... 5-12 ROTOR SYSTEM K;NEMA IC COU?U]N!G ......................... 5-13 Pitch-lag Instability ................................................ 5-1; Pitch -flap Instability .............................................. 5-14 Flap-1iit Instability ............................. .................. 5-14 ROTOR SYSTEM DYNAMICS ....................................... 5-!6 OSCILLATORY LOADING 01- ROTOR BLADES .................... 53-16 Hyotheicator LoadDsigcdinCofsid oratibns..............................5-16 Osilltoypohtiad DsesaigConsidrations V ratoy.Load....................51 .. Rotor Oscitlatory Load Calculation....................... .......... 5-17 Drawing Pomid Phase ....................... .. .................. 5-17 Flight rests ..................................................... 5-18
5-4.1.2.4
.
_ ................
5-18 5-19
5-2 1 5-22 5-22 5-23 5-23 5-23 5-23
:
1 '
5-23 5-23 5-23 5-23 5-24 5-24 5-24 5-24 5-25 5-2ý 5-26 5-1) 5-27 5-27 5-27 5-27 5-29 5-29 5-29 5-30 5-30 5-30 5-30
.
"
AMCP 706-22 "1AII: kI F OFl
()N"t I-NS t( outinu'd i
Piaragraph
5-5.2.1. 5-5.2.1.2 5-5.2.1 3 5-5.2.2 5-5.2.3 5-5.2.3.1 5-5.2.3.2 5-5.2.4 5-5.2.5 5-5.2.6 5-5.3 5-5.4 5-5A.1 5-5.4.1.1 5-5.4.1.2 5-5.4.2 5-5,,4,3 5-6
5-6 IF
5[ -t,.t 5-i. 1 2 5-6.1.3
5-6.2 5-11! 2. 1 5-6.2. I .I 5-6.2.1.2 5-6.2.1.3 5-6.2.!.4 5-6.2.1.5
5-6.2.1.6
Ig
C_')hndrical Roller Brarinis ........................................... Tapered Roller Bearings .............................................. Angular Contact Ball Bearings ........................................ Tcflon F-abric Bcaring. ................................................ Flexing Elk nent.' ...................................................... Tension-torsion Strap Assemblies .................................... W ire Tic-bar Assem blies ............................................. Flastom criý" Bcarings ..................... ............................ Lag D)ampcrs, Lead-lag Stops ......................................... Droop and Flap Stops and Rcsti ainers ................................. CONTROL SYSTEM CONSIDERATIONS ............................ BLA D E FO LD ING .................................................... D esign Requirem ents .................................................. M anual Blade rolding ........................................... .... Pow er Blade Foiding ................................................. O perational Requirem ents .............................................. System Safet) Considerations .............. ............................
R OTO R BLA D ES ........................................................
N q, I
5-6.2.4
5-6.2.5 5-6.2.6 5-6.2.7 5-6.2.8 5-6.3 5-6.3.1 5-6.3.2 5-6.3.3 5-6.4 5-7 5-7.1 5-7.2 5-7.2.1 5-7.2.2 vWi
.................
... . .............................
BLADE CONSTRUCTION ........................................ S p a r ................................... .............................. H ollow Extrusion .................................................... Solid Extrusion ..................... ................................ Formed Sheet Metal .................................................. R ound S,ecl T ube ...... ............ ................................ M olded Reinforced Plastic ............................................ le I UvI I .........................
.......................
.
5-37
5-38 5-38 5-39"•
(
1 : ':'
, ':
5-41 5-4 1 5-41 5-41i-_ 5-42 5-42
, S,
5-42
I ....... .
C ontinuous Skins .................................................... Segm ented Skins ............................... ................ .... W raparound Skins ................................................... ............................... Root End Retentions ................ Tip Closures and Hardware .....
5-37
5-42
Form ed M etal Tube .................................................. .2...rt
5-6.2.2.1 5-6.2.2.2 5-6.2.2.3 5-6.2.3
. ......
"Tw ist ................................................................. Planfurni " apt~r ... ..... .............................................. A irfoil C ross Section ................... ...............................
5-31 5-31 5-31 5-31 5-32 5-32 5-32 5-32 5-34 5-34 5-35 5-35 5-35 5-35 5-3t 5-36 5-36
......................................
T rim T abs .................... ....................................... T uning W eights .............. ........................................ D esign Requirem ents .................................................. Tooling and Quality Control P equirements ............................. BLADE BALANCE AND TRACK ..................................... Effect of D esign ............... ....................................... Com ponent Lim it W eights ............................................. 1 rack ................................................................. ROTOR BLADE MATERIALS ........................................ ROTOR SYSTEM FATiGUE LIVES .................................... G EN E R A L ........... ................................................ ENDURANCE LIMIT TESTING ...................................... G encral ............ .................................................. N onm etals ........ .................... ..........................
5-43 5-4 3 544
5-44 5-44
5-44 5-4 5 5-4 5 5-45 5-46 5-4 6 5.47 5-4 9 5-50 5-53 5-53 5-54 5-54 5-55
'
.-
,
1
F~ AMCP ""1 ABI IV!
(
706-20
%'11 iN SI( onfilud
PIr.iagraph
,
...............................................
5-57
5-8
PROP
II. RS .......
.... ..
. ...............
5-H. I 5-8.2 5-3.2. 5-8.2 2 5-8.2.3 5-8.2.4 5-8.2.5 5-8.3 5-8.3.1 5-,.3. 5-8.3.1.2 5-8.3.1.3 5-S.3.2 5-8.3.2.1 5-8.3.2. 1.1 -
GI NL.RAI ............... .................................... PROPELLER SYSI ENI )YNAMICS .................................. Vibrator) L.oads ...................................................... . Critical Speeds and Resiporisc .................. ....................... ....... ............................... G usis and M aneuvers ...... Stall F luttcr ....................................... .................. Prop Iller Roughness ...... ....... ............................ .... PROPI-I.iI-IR HUBS, ACTUATORS, AND CONI ROLS ............... Propeller Barrel and hladc Rctention . ................................ B arrel L.oading ....................................................... Louding D efinition .......................... ....................... Barrel Structural Tests ........................................... Propclcr Actuators and Controls ...................................... C ontrol Configurations ................... ........................... Constant-speed G overnors .......................................... .. ........................................ BnaContr,..
5-57 5-s5 5-57 5-60 5-62 5-63 564 5-65 5 65 5-65 5-66 5-66 5-66 5-66 5-66 5-67
5-8.3.2.3 5-8.3.2.4 -5-8.3.2.5 5-8.4 5-8.4.1 5-8.4.2 5-8.4.2.1 5-8.4.2.2 5-8.4.2.3 5-8.4.3 5-8.4.4 5-9. 4.4.1 5-8.4.4.2
5-67 5-68 5-68 5-18 5-68 5-70 5-70 5-7.2 5-73 5-73 5-7. 5-74 5-75
5-8.4.4.4 5-8.5 5-8.5.1 5- .. .S
Au xiliary Functio s .................................................. ......................................... Control Performance C ontrol R eliability ................................ .................. PROPELLER BLADES ......... ................................ Blade Geom etry ...................................................... Blade C onstruction .................................................... ade Constructior .......................................... Types of BM M anufacturing Processes and Tooling ................................. Q uality C ontrol ........................................... .......... Blade and Propeller Balance ....... ....... ........................... Bvide M aterials ................................. ..................... tl low B lades ....................................................... C om posite M aterials ................................................. ................. F iller M aterial ..................................... Structural Adhesives ....................... ......................... PROPELLER BLADE FATIGUE LIVI:S ............................. Enduranct Limit and O0her Structural T',:sting .......................... ... .......................... ........... . T i A TIst C... I.
5-75 5-75 5-76 5.7
5-8.5.1.2 5-8.5.2
F ull-scale Tests .................................... .................. tigu Life Dctcrminat on ................ Flight Loads Test D-aon
5-76 5-76
IO p erpretation of ResutsI ..............................................
5-77
5-9.4.1 5-9.4.2
Tractor C onfiguration ................................................. ................... Pusher C onfiguration ...............................
5-79 5-79
5-9.4.3
O perational Considerations ............................................ D irection of R otation ..................................................
5-79 5-79
5-,.4.4.3
5-H.5.2. I
5-8.5.2.2
)•
5-56 5-56
Structurai M em bcr I. .............. 5-7.2.4 Dctermination of Fatigue .ifc
5-8.3.2.2 --
...................
5-7.2.3
5-9 5-9.1 5-9.2 5-9.3 ,5-9.4
S5-9.4.4
Itydraulic System
..........................................
A ircraft T ests ........................
............ . ........
. ........
. .......
............................ AN -IITORQ U E ROTORS ................... . ........... ........... .................... G E N E R A L ............... .......... TYPICAL. ANTiIORQU)E ROTORS ........................ TAIL ROTOR DESIGN REQUIR!M ENrS ........................... .... ....... INSTAILLATION CONS!IDERATIONS ....................
5-67
5-75
5-76
5-7"7 5-77 5-78 5-78 5-79
.,
AM^P 706-202
f 1 AlI.t OE ( Oi11 -N'i S Iollhunitvd (
Paragraph
Pap" Engint E.xhaust
5-9 4 5
5-9.5 5-9.5.1 5.9.5.2 5-9.5.3
"5-9.5.4 5-9.5.5 5-9.6
5-9.7 5-9.7.1 5-9.7.2 5i-9.7.3 5-".7.4 5-9.7.5
* I
-. -_.
6-0 6-i 6-1.1 6-1.I.I 6-1.1.2 6- .1.3 6-1.2 6-1.3 6-2 6-2.1 6-2. 1.1 6-2.1.2 6-2.1.2.! 6-2.1.2.2 6-2.1.2.3 6-2.1.2.4 6.2.1.3 6-2.1.4 6-2.1.5 6.2.2 6-2.3 6-3 6-3.1 6-3.1 .1 6-3.1.2 6-3.1.3 6-3.1.4 6-3.1.5 6-3.1.6 6-3.1.7 6-3.2 6-3.2.1
.A
......................................................
. 80
TAIL ROTOR DESIGN PARAMETERS .............................. Tail Rotor D isk I oading ............................................... T ail Rotor Tip Speed .................................................. ....................... .......... Bladc Nuinocr and Solidity ..... T wist ............. ................................................... 0lade Airfoil t........................................... ...... TAIL ROTOR PI RIORMANC ....... ......... STRUCTURAL CONSID'ERATIONS ............. ............. Struclural D ynam ics ............................ ..................... . Structural Loading ....................... ....................... .. .. Blade Structural Analysis ............. .. .. ......................... A .- uelasticit% ............ ............................................ Flutter and Divergence .......................................... R EFER EN C ES .. ............................... ..... .................
5-89 5-80 5-80 55-8I 5-RI 5.2 5-82 5.82 5-82 5-83 5-83 5-83 5-83
(IAlv IKR 6 FI.V;HT CONTROL S. 1•SVS,,iLIST O t SYM BO LS .................................................... GENRA, .L................................................... ......... ........... D ESIG N M ETIIO ). ............... I Point of Dcpartuie ............... .............................. Mission Requirements and Fligh" Envelope .......................... Basic ttclicoptc" D ata ............................. ........... ...... ANALYTICALTOOLS .......................................... SIMULATION AND TESTING .................................... STABILITY SPECIFICATIONS ........................................ C RITERIA AND METHOD OI'.\V'AI YSIS ........................... Control Power and Damping .......... .............. ............... C haracteristic R oots ................................................... R om Plot, .............................................. .......... Modes and Required Damping ..................................... Inherent Airfram e Stability . ....................... ................. Variation of Para;m eters ... ................... ...................... T ype of C ontrol ............................. . . ..................... Transient Response .. ......... ............. ................. Other Factors ................................................ AIUTOROTATION INTRY . ..................................... SYSTEM FAI URLS ............................ ............ ...... STA BILI FY AUGMENTATION S%Si'i-MS .......................... .. GENERAI ..................... .......... ............ ..... B ell Stabiliter Bar ............... ..................................... H Servo GRotor ............. ......... . ............ ................... ......... M iller echanical -,ro .............. ...................
6.6 1 6-i 6-I 6-1 6-2 6 -6-2 6.2 6-2 6-2 6-i 6-3 6-4 6-4 6-6 6-6 6-6 6-6 6-7 6-8 6-8 6-9 6-9 6-9 6-9 6-9
Lockheed C ontrol Gyro ............................ .................. Elcctrohydraulic SA S ...................... ......................... . Fluidic and Hydrofluidie SAS ....... ................................. Flapping M oment Feedback ....... ......... ..................... . CRITERIA FOR SELECTION .................................... A ugrm entation Requirem ents ................................. .........
6-10 6.13 6-10 6.10 6-10 6-10
(
r
AMCP 706.202 TABLE OF CONTENTS (Continued) Pagt
Paragraph
6-3.2.2
H elicopter Size ........................................................
6-10
6-3.2.3 6-3.2.4 6-3.2.5
Type of Rotor System ................................................. Helicopter Configuration .............................. ............... Suppression of Structural and Rotor Mode Responses, Vibrations. o; G usts ............................................................. SAS RELIABILITY .................................................... ... ..................................................... S afety ..... SA S Fa:lures .......................................................... Fail-safe Principles .................................................... Battle Damage, Vuhnerability ........................................... .......................................... C OST ....................... Developm ent Cost ... ................................................. Production C ost ....................................................... M aintenance Cost ..................................................... TECHNICAL DEVELOPMENT PLAN ................................ PILO T EFFO RT ............................ ............................ CRITERIA FOR POWER CONTROLS ................................ Control Forces .............................................. Vibration lFeelback ............................................ .......................................... K in .inatic Effects ........ Control Stiffness ........... .......................................... HANDLING QUALItY SPECIFICATION ............................ HUM AN FACTORS .............. ................................... Control Force C ues ............. ......... .................... ....... Developm ental Test ................................. .................
6- 1 6-11
6-3.3
S6--?.r`.l 6-3.3.2 6-3.3.3 6-3.3.4 6-3.4 6-3.4.1 6-3.4.2 6-3.4.3 6-3.5 6-4 6-4.1 6-4.1.1 6_6-4.1.2
S ,. A
.
) ,
6-4.1.3 6-4.1.4
6-4.2 6-4.3 6-4.3. 1 6-4.3.2
AUTOMATIC CONTROL INTERFACES .............................
6-4.4
VU LN ERA BILITY .................................................... R ELIA BILIT Y ........................................................ M EC H A N ISM S ......................................................... YSTEMS .............................................. ROTATING a Design Factors ........................................................ T est R esults ........................................................... Bench Tests .......................................................... .................. . Test Loads .............................. Instrur"-,ntation ..................................................... Quantity and Selection of Specimens ............................. lnterpretation of D ata ............................................... Flight T ests ........................................................ Required Instrum entation ........................................... Flight Cond1 .ions ........................................... NONROTATING SYSTEM ........................................... Pilot's Controls to Power Actuator ......................... .......... Power Actuator to the Swashplate ...................................... T R IM SYST EM S ..................................... ................ D isconnectT ri ...................................................... C ontinuous rrim ................................ ..................... Parallel and Series T rim ................................................ SYSTEM DEVELOPMENT .............................................. ....................................... G E N E R A l . .................... MATHEMATICAL MODEL IMPROVEMENT ......................
6-4.5 6-4.6 6-5 6-5.!i 6-5.1.2 6-5.1.2.11 o-5.l.2.1.1 6-5.1.2.1.2 6-5.1.2. 1..1 6-5.1.2.1.4 6-5.1.2.2 6-5.1.2.2.1 6.5.1.2.2.2 6-5.2 6-5.2. 1 6-5.2.2 6-5.3 6-5.3.1 6-5.3.2 "• 6.5.3.3 6-6 6-6.1
6-11 6-12 6-12 6-12 6-13 6-13 6-13 6-13 6-13 6-13 6-13 6-14 6-14 6-14
6-1A 6-14 6-15 6-15 6-17 6-17 6-17
6-17 6-18 6-18 6-18 6-18 6-ig 6-21 6.21 6-.21 6-21 622 6-22 6-22 6-2", 622 6-22 6-23 6-24 6-25 6-25 6-25 6-26 6-26 6-26 6-26 xi
-- m.'ll .-.
~
Il
- ...- •t&
'.1
*
0.
I~Ai • ae&ih
"
,,l.a.&
Mkh~msl,.
.-.-...-.-..
.
.
~ --
.
-. i
AMCP 7QW5V2 TABIIV OF ('ONI 1.N*1S t CoEn~inut i) Paragraph 6-6.2.1 6-6.2.2 6-6.3 6-6,4 6-6.5
Pagc ..................... Wind Tunnel Test............................ .. I......................... Hardware Bench Tests ................. GROUND-BASED PILOTED FLIGHT SIMULAT ION ................ FLIGHT TESTS................................................... DESIGN REVIEW ................................................ REFERENCES .. ..................................................
CHAPTEFR 7 ELEC'TRICAL SUBSYSTEM DESIGN 7-0 LIST OF SYMBOLS................................................. 7-I INTRODUc-riON.................... ....... ....................... 7-1.1 GENERAL....................................................... 7-1.2 SYSTEM CHARACTERISTICS..................................... 7-1.3 LOAD ANALYSIS ................................................ 7-1.4 LOAD ANALYSIS PREPARATION ................................. 7-1.5 MANUAL FCRMAT .............................................. 7-1.6 AUTOMATED rORMAT ............................. ........... 7.1.7 SUMMARY...................................................... 7-2 GENERATORS AND MOTORS ...................................... 7-2.1IC GEEAI........................... .:-. ................ 7-2.2 AC GENERATORS (ALTERNATORS).............................. 7221Eiectrical Design................................................. 7-2.2.2 Mechanical Dcsign ................................................ 7-2.2.3 Cooling........................................................ 7-2.2.4 Application Checklist .............................................. 7-2.2.5 V aria ble-freqqtency AC Generators................................... 7-2.3 srA RTER/G EN ERATO RS, DC G IN ER ATORS, AND STA RTERS 7-2.3.1 Starter/Generator ...................................... ......... 1-2.3.2 DC Geuerators....................... ............................ 7-2.3.3 DC Starters ...................................................... Boost Starting System .............................................. 7-2,3.4 1-2.4
7-2.5 7-2.5.1 7.2.5.2 7-3 7-3.1 7-3.2 7-3.3 7-3.4 7-3.5 7-4 7-4A 7-4.1.1 7-4.1 2 7-4.1.3 7-4.2 7-5 7-5.1 7-5.2
ELECTICAL MOTORS............................
....................
ELECTRICAL SYSTEM CONVERSION............................. AC to DC Converters ............................................. DC toAC Converters .............................................. ...................................... BATI'LRIES ................. Bm rERY CHARACTERISTICS ................................... ~ GENERATOR CONTROL BATTERY CHARGING................... .......... UTILIZATION LOAD ANALYSIS........................ HEAVY CURENT STARTING REQUIREMENTS .................... MAINTENANCE ......... ..................................... VOLTAGE REGULATION AND REVERSE CURRENT RELAY ... DC VOLTAGE REGULATION ..................................... Voltage Regulator ....................... ............ ............ Reverse Current Relays........................................... Overvoltage Relays ..................................... .......... AC VOLTAGE REGULATION ..................................... OVERLOAD PROTECTION ......................................... GENERAL....................................................... OVERLOAD PROTECTION DF-VICES ............................ .
.26 6-27
6-27 6-2ý, 6-28 6-28
7-1 7-1 7-1 7-1 7-2 7-2 7-3 7-3 7.4 7-4 7-4 7-6 7-6 7-6 7-7 7.8 7-8 7-9 7-9 7-10 7-11 7-1l 7-13
7-14 7-14 7-15 7-15t 715 7-15 716 7-17 7-18 7-11... 7-18 7-18 7K 7-19 7-19 7-19 7-19 7-20
l
A
F .AMCP
706-202 1IABI 1. 01 (()N'II:NI I,(
nttinued)
Paragraph 7-5.2. 1 7-5.2.1.1 7-5.2.1.2 7-5 2.2 7-5.2.3
"7-5.2.4 7-5.3 7-6 7-6.1 7-6.2 7-6.3 7-6.4 7-6.5
"7-6.5.1 7-6.5.2 7-6.5.3 7-6.5.3.1 7-6.5.3.2 -6.5.3.3 7-7 7-7. i ?7-7.2 7-7.3 '• 7-7.4 7-7.5 7-7.6 7-7.7 7-7.8
"7-8
7-8.1I 7-8.1. 7-8.1.1.1
7-8.1. 1.2
7-8.1.1.3 7-F8..1.4 7-8.1.1.5 7-8.1.1.6 7-8.1.1 .7 7- . Q.2 7-8.2 7-8.2.1 7-8.2.2 7-9 7-9.1
"1-9.2 7-9.3 7-9.4
Circuit Breakers ................................................... 7-20 Therm al C ircuit Breakers ............................................. 7-20 7-20 ............... M ars.czic Circuit Breakers .......................... Remote Control Circuit Breakers ....................................... 7-20 C urrent Sensors ...................................................... . 7.20 F uses .............................................................. . . 7-20 OVERLOAD PROTECTION APPLICATION ......................... 7-21 ELECTROMAGNETIC INTERFERENCE (EMI/EMC) ................. 7-21 G E N E R AL ............................................. ............... 1-21 ACCEPTABILITY REQUIREMENTS ................................. 7-21 INTERFERENCE SPECIFICATIONS ................................ 7-21 !NTERFERENCE SOURCES ........................................ 7-21 INTERFERENCE SUPPRESSION ................................... 7-22 Interference-free Components ......................................... 7-22 Equipment Isolation and Cable F outing ................................ 7.22 Source Suppression and Susceptibility Reduction ..................... 7-23 G rounding and Bonding .............................................. 7-23 Shielding ............................................................ 7 -23 F ilters ............................................................... 7-24 ELECTRICAL SYSTEM INSTALLATION ............................. 7-24 -'% N E R A i .. ............................................................ h IE EQUIPMENT INSTALLATION ....................................... 7-24 ELECTRICAL WIRE BUNDLES ....................................... 7-25 TERMINAL STRIP INSTALLATION .............................. 7.25 ENGINE COMPARTMENT WIRING ................................. 7-26 DOOR HINGE WIRE BUNDLE ROUTING ........................... 7-26 WIRING TO MOVING COMPONENTS .............................. 7-26 BATTERY INSTALLATION .......................................... 727 CO M PO N EN TS .............................. ......................... 7-27 7-27 W IR E ................................................................. W ire Insulating M aterials ...... ....................................... 7-27 Polyethylene ......................................................... 7-27 Polyvinylchiorid¢
. . . . . . . .. . . . . . . . . .. . . . . . . . .
Fluorinated Ethylene Propylen ................................. Polychlorotrifluoroethylene ............... ..................... Polybexamethylene-adipamide ............. .......................... Tetrafluorocthylene .................................................. Dimethyl-siloxanc Polymer M ilitary W ire Specifications ........................................... F I-f IN G . ............................................................ Term it,al Strips ........................................................ Connectors ........................................................ . . LIGHTNING AND STATIC ELECTRICITY ............................ G EN ER A L ............................................................ LIGHTNING PROTECTION FOR ELECTRONIC SUBSYSTEMS .... STATIC ELECTRICITY ............................................... LIGHTNING AND STATIC ELECTRICITY SPECIFICATIONS ...... R EFER EN C E ...........................................................
7-27
7-28 7-28 7-28 7-28 7-28 7-28 7-29 7-29 7-29 7-29 7-29 7-30 7-31 7-32 7-32
xiii
.
-
-:
r• =.-''"
TABLE OF CONTENTS (Conlinued) Paragraph
8-11 I.;.. 8- L2 8-1.3 8-1.4 8-2 8-2.1 8-2.2 8-2.3 8-3 8-3.1 8-3.2 8-3.3 8-3.3.1 8-3.3.2 8-3.3.3 8-3.3.4 8-3.3.5 8-3.3.6 8-3.3.7 8-3.4 9-3.5 I.8-3.6
8-4 8-4.1 8-4.2 8-4.3 8-4.4
Pagc CHAPTER 8 AVIONIC SUBSYSTEMS D~ESIGN INTRODUCTION ................................................. 8GENERAL.................................................... 8i ELECTROMAGNETIC COMPATIBILITY PaOGRAM .. ............. &I DESIGN CONSIDERATIONS..................................... 8-2 ENVIRO1NMENTAM. ASPECTS.................................... 8-2 COMMUNICATION EQUIPMENT ........................ ......... 8-3 GENERAL ..................................................... 8-3 MICRt)PIONE-HEADSET ....................................... 8-4 INTERCOMMUNICATION SELECTOR BOX......................-., NAVIGATIONAL EQUIPMENT;................................... .8-4 GENERAL ..................................................... b-4 TERMINAL MANEUVERING EQUIPMENT ....................... 8-5 EN ROUTE NAVIGATION EQUIPMENT........................ h5 Automatic Direction Finder (A.DF) ................... ...... ....... 8-5 Distance-measuring Equipment (DMW) ................ I............. 8-5 Tactical Air Navigation (TACAN).................................. 845 Lona-ranize Navistation (LORAN).................................. t-6 Compasses ..................................................... 8-6L Doppler Navigation Systems ....................................... 8-6 Inertial Navigation Systems ................................. ...... -E; 6 INTERDICTION EQUIPMENT ................................... 8_/ 8-7 1.0W-LIGHT-LEVEL NAVIGATIONAL EQUIPMENT ............... STATION-KEEPING EQUIPMENT .............. ............... 9-7 FIRE CO1ý'TROL EQUIPMENT ......................... .......... 8GENEt. AL................................. .................... 8-7 INSTALLATION .................................. .. ........... 8-8 SIGHTING STATION ............................................ 8-8 SENSORS ...................................................... 8-8 ..
k
~
.-
T~rvDW
O At C'~k
8-4.6 8-4.6. 1 8-4.6.2 8-4.6.3 8-4.7 8-5 8-5.1 8-5.2 8-5.3 8-5.3.1 8-5.3.2 8-5.3.3 8-5.3.4 8-5.3.5 8-5.3.6
IX
..
..
. .
. .
. .
.
II. .
.
. .
.
. .
.
.
. .
.
.
.
.
.
. . .
FIRE CONTROL ACCURACY................ I.................... Inertial Stabilization.............................................. Fire Control Datum Planc ..... ................................... Harmonization .................................................. COMPONENT LOCATION......................... .............. ANTENNAS ...................................................... GENERAL ............................. ..... .............. .. ANTENNA DEVELOPMENT..................... ................ LOCATION AND INSTALLATION OF ANI ENNAS ......... Communication Antenna Considierations ............. .... Low Frequency (-1) ........................ .............. ...... H igh Frequency (H F) ý...... .................................... Very High Fre~quency (VHF).......... .................... ... .... Ultra High Ficqucncy (UHF)........... ..... ............ ........ SpeCi3l Purpose ................................................. REFERENCES........................................... .........
r
I
)Lk
8-9 8-9 8-9 8-10 8-10 8-10 U-11 912 f 3 l-l3 : 8-14 8-14 8-;4 g9.14
tA
[_,0
TABLE OF CONTENTS (Continued) ParRAUapC *LISTOF
N
.-
9-2.1 9-2.1.1 -9-2.1.2 9-2.2 9-2.2.1 9-2.2.2 9-2.2.3 9-2.2.4 9-2.2.5 9-2.3 92.. 9-2.ý.2 9-2.3.3nu 9-2.4 9-2.5 9-2.6 9-2.6.1 9.2.6.1.1 9-2.6.1.2
I9-2.6.1.3 9-2-6.1.4 ý92.6.2 '92.. 92.6.4. V-ZAA.1
4\
AND:NEUATI
SUBSYSTEMS DESIGN
SYMBOLS ...............................................
9-1
FLIGHT CONTROL POWER SYSTEMS.......................... Central Hydraui2c Sysiemn......................................... Flight Control Subsystems........................................ UTILITY HYDRAULIC SYSTEMS ................................ Engine-starting Subsystems........................................ Cargo Door and Ramp System ..................................... Cargo and Pnrsoniicl Hoist ........................................ Rotor Brake ..................................... .............. Wheel Brakes................................................... H-YDRAULIC SYSTEM RELIABILITY ......................... ... Flig ht Control Redundancy........................................9Utility System Redundancy........................................ R elinhilitv Asnacte ....................... 41 YDRAULIC SYSTEM STfRENGTH- CONSIDERATIONS............ H-YDRAULIC SYSTEM TEMPERATURE CONSIDERATIONS .... HYDRAULIC SYSTEM DESIGN ......................... ........ Sun'vability. Reliability, and Safety Trade-offt........................ Reservoir Leve! Sensing ................ .........................
91 9-1 .9-2 9-2 9-2 9-3 9-4 9-4 9-59-S
Systemt Switching Concepis
..........
..
-- 8
......
Return Pressure Sensing ....................................... .. Switching and Return Pressure Sensing..................... ...... Operating Pressure Considerations.................................. Selection of Fluid Medium ............... ......................... Filtration (Contamination)........................................ Fluiud
Fiiirutiun. ..............................................
9-5 9-5 9-6 9-6 9-7 9-7 9-8
.
1.. K 1
9-9 9-9 9-10 9-10 9-104
9-!0
9-Il 9-2.6.4.2Ground Operation Filtration ..................................... 9-2 6.4.3 Filtration Lcvel ................................................ 9-11 9.2.6.4.4 iex~rnal Contamination................... ...................... 9-11 9-2.6.5 Fittings ....................................................... 9-1 i 9-2.6.6 Dynamic Fluid Connections.................................. ..... 9-12 9-2.6.7 Peak Power Leve~ls............................................. 9 9-2.6.8 A1PU ant' Engine Starting ................................... ... . 9-2.6.9 Syskem I-eat Rejection Characteristics ................. ........ ... 9-13 9-2.6. 10 System Analysis........................... ...................... 9-13 9-2.7 HYDRAULIC COMPONENT DESIGN AND SEL ECTIO ... ......... ý~ 9_1I ............. Actuators ... L...................................... 9-2.7.1 9-2.7.1.1 Rip-stop Protection ............................................. 9-0 9-?.7. i.2Endurance Testing Requirements.................................. 9.4g 9-2.7.1.3 Seal Alternatives ....................... ........ ............. 9 92.7.1.4 Materials and Stres!; Corusidci tions ................ ............. .. 9-16 9-2.7.1.5 General Requirements...................................... ..... 9- 1f 9-2.7.2 Hydraulic Pumps;.............................. .. ... ............ ~* ~ k9-2.7.3 Accumulators................................................. 92.. Reservoirs ............ L................... ....... ........ ..... 4-20 ....... .. ........ ... e-2c 9.2L-7.5 Pressure Relicf............................ Nx
4
V
'
AMCP 706-202
,T".I Tmo0 (ONvl I-"N'IS ,
qmtinued u
Paragraph 9-2.7.6 9-2.7.7
"9-2.7.8 9-2.7.9
9-2.7.11 :9-2.7.13
9-2.7.12 9. .7. ,z
9-2.7.14 9-2.8 9-2.8.1 9-2.8.2
"9-2.8.3 9-2.8.4
9-2.8.5 9-2.9 9-2.9.1 9-2.9.2 9-2.9.3
"9-2.9.4
,9-2*.
:
9-20 921 9-21 922 9 .22
Control Sclttor Valves ......................... ................. R estrictors ............ ............................................ S~parauc Secvos a a c e v , ............................. ............ .............
9-22 9-24 9-25;
... ..
9 -25.
AIlovable External Leakage ........................................... HYDRAULIC SYSTEM INSTAILLATON ........................... U se of Hoses and Sw ivels .............................................. Ma-ntrtcance Access ............................................ Hard Vetsus Soft !h.s,stllativns ......................................... Component M ou zingnConctnts .......................................
9-25 9-25 9-25 9-25 9-26 9-26
Miscellaneous Instahiation Conside-ati 1 ... ......................... MISCELLANEOUS DESIGN CRI IERIA ............................. Actuators and Assoziated Equipment Design ...... ................... Brake Design .......................................................... Control System Design ................................................ Electrical Design ................................. ....... ............ , ..........................................................
9-27 9.27 9.27 9-28 9-29 9-29
9-2.9.6
Fittings Design . ......................................................
9-;2c).D 9-30
9-2.9.7 9-2.9.8 9-2.9. 9-2.9.10 9.2.9.11
G age and Indicator Design ............. ............................... H osP D esign .......................................................... s .......................................................... R eservoir Design ...................................................... Valve D esign ..........................................................
9-30 9-30 9 -30 9-31 9-31
9-2.9.12
Lubrication ........................................................... PNEUMATIC SYSTEMS.....G ..................................... PNEUMATIC SYSTEM DESIGN...................................... System A nalysis. ... .... ...................... .....................
9-32 9-32 9-32 9-32
"9-3 9-3.1 9-3.1.1
9-3.1.2
System Redundancy
.. ....................... .....................
9-33
"9-3.2.1
COMPONENT DESIGN ......................................... A ir C om pressors ............ ........................................ Positive D isplacem ent ................................................ Dynamic Displacement ..........................................
9-33 9-33 9-333 9-34
9-3.2.2 9-3.2.3 9-3.2.4
Compressed Air Supply System Selection and Operation ................ Moisture Separators ................................................... Dehydrators ..........................................................
9-34 9-34 9-35
9-3.2.5 9-3.2.6 9-3.2.6.1
F ilters ................................................................ Valves ...................................................... Check Valves ................................................
9-35 9-36 9-36
9-3.2 9-3.2.1.1 9-3.2.1.2P
9-3.2.6.2 9.3.2.6.3 -9-3.2.6.4 9-3.2.6.5 v -3.2.7
,.'.., •~
Pressure Regulation ................................................. Filters .................... .................................. C heck Valves ....................... .......................... ....... Pressure Switches .............................................. PressureTr~nsm !92.7.10 itters ............................ ................ ...
R elief Valves ......................................................... Pressure-reducing Valves .................................... Pressure R egulators .................................................. Directional Control Valves ................................... u ........................................................
9-3.2.8
A ir Storage Bottles ....................................................
9-3.2.9 9-3.2.9.I
Subsystem Com ponerts ............................................... Actuators ............................................................
9-37 9-38 9-38 9-38 9-41 9-42
9-43 9-43
" .
Li
AM0CP706-202 "IAB1.4 01 ()%'
IN' I
SI ( ontinucd)i
Paragraph 9-3.2 9.2
9-3.2.9.3 9,3.2.9.4 9-3.3 9-3.4 9-3.4 .1 9-3.4.2 9-3.4.3 9-3.4.4 9-3.4.5 9-3.4.6
.
10-1 10-2 10-2.1 10-2.2 10-2.3 !0-2.4
-
10-2.4.1 10-2.4.2 10-2.4.3
10-3 10-3.1 10-3.2 10-3.3 10-3.4 10-3.5 10-3.6 10-4 i-4.i
10-4.2 10-4.3 IO.5 10-5.1 10-5.2 10-6 10-6.1 10-6.2 10-6.2.1 10-6.2.2 10-6.2.3 10-6.3 10-7 10-7.1 )10-7.2
10-7.3
Page Brake V alves ..............
..........................................
9-44
Pneum atic Fu ,cs ..................................................... Q uick-disconnect. ............... ............................. ...... PNEUMATIC SYSTEM INSTALLATION AND QUALIFICATION ... PITOT-STATIC SUBSYSTEM DESIGN ............................... A ltimr et rs ................................... ........................ Rate-of-clim b Indicator ............................................... A irspccd Indicators .................................................... Total-pressure Sources ................................................. Static Pressure Sources ................................................ Pitot-static T ubes ...................................................... R EFE R EN C ES ..........................................................
9-44 9-44 9-44 945 9-4 5 9-46 9-46 9-47 9-47 9-4 8 9-48
CIIAPTER 10 INSTR I EMiNTATION SIUBSYSTEM I)ESIGN IN T RO D U C TIO N ............................. ........................ INSTRUMENTATION LIGHTING REQUIREMENTS ................. G EN E R A L ............................................................ LIGHTING INTENSITY CONTROL .................................. LOW INTENSITY READABILITY ....................................
10-1 10-1 10-1 10-1 10-2
WARNINGC. CAUTION, AND ADVISOR Y S!ONALS
................
W arning Signals ....................................................... C aution Signals ... ................................. .................. A dvisory Lights .......................................................
10-2 10-2 10-3
FLIG HT INSTRUM ENTS . ............................................. ( EN ER A L ............... ............................................ AIRSPEED INDICATORS ......................... .................. A LT IM ET ER S ........................................................ TURN-AND-BANK INDICATORS ................................... ATTITUDE INDICATOR ............................................. RATE-OF-CLIMB INDICATORS ..................................... NAVIGATIONAL INSTRUMENTATION .............................. ENEAL...........................................................
1.7
10-3 10-3 10-3 10-3 10-3 10-3 10-5 10-5
....
TYPESOF INSTRUMENTS ........................................... M A P DISPLA YS ...................................................... HELICOPTER SUBSYSTEM INSTRUMENTATION .................... G EN E R A L ................... ............................ ........... INSTRUMENTATION REQUIRED .................................. WEAPON SYSTEM INSTRUMENTATION ............................. G EN ER A L ............................................................ DESIGN REQUIREMENTS ........................................... Arming, Fuzing, and Suspension and Release Control Design ............ Human Fa -tors Considerations ......................................... Indicator Dcsign ...................................................... WEAPON SELECTION CONTROLLLR/PROGRAMMER ........... TYPES Of INSTRUMENT .............................................. IN STA LLATIO N ...................................................... V IBR A T IO N .......................................................... ACCESSIBILITY AND MAINTENANCE ............................. R EFER EN C ES ............................................ .............
10-5 10-7 10-7 10-7 10-7 10-7 10-7 10-8 10-" 10-8 10-8 10-9 10-9 10-10 10-10 10-10 10- 10 xvii
TABLE OF CONTENTS (-]onitnuvdi Paragiaph
I -f0 I-oI 11-2 11-2.1 11-2.2 1a-2.3 11-2.4 11-2.5 11-2.6 11-2.7 11-3 11-3.1 11-3.2 11-3.3 11-3.4 11-3.5 11-3.6 i i-3.7 11-4 11-4.1 11-4.2 I1-5 11-6 11-7 I-8 11-8.1 I1-8.2
12-0 12-1 12-1.1 12-1.1.1 12-1.1.2 12-1.1.2 .1 12-1.1.2 .2 12-1.1.2.3 12-1.1.2.4 12-12 12-1.2 ,1 12-1.2.2 12-1.2.3 12-1.3 12-1.3.. 12-1.3.2 12-1 3.3
"xViii
Page CHAPTER i i AIRFRAME STRUCTURAL DESIGN LIST O F SYM BO LS ..................................................... INTRO DUCTIO N ....................................................... I)ESIGN CONSIDERATIONS ........................................... W EIG H T .............................................................. SURFACE SMOOTHNESS ............................................ STIFFNESS AND RUGGEDNESS .................................... FATIGUE SENSITIVITY ............................................. C O ST .................................................................. M A T ER IA LS .......................................................... SU RVIVA BILITY .................................................... DESIGN AND CONSTRUCTION ....................................... FITTINGS .................................................. SU PPO RT S ............................................................ FRA M ES ............................................................. BU LK H EA DS ......................................................... SKIN SUBSYSTEM S .................................................. CORROSION PROTECTION ........................................ FiEC.TRWiAL BLONDiNG ................................................ CARGO COMPARTMENT .............................................. STA T IC LO A D S ....................................................... C RA SH LO A D S ...................... ................................ TRANSPARENT AREAS ................................................ D EV ELO PM EN T ............ ........ .................................. M A N U FACTU R E ................. .......................... .......... SU BSTA N TIATIO N ..................................................... A N A L Y SIS .............................................................. T EST IN G ................................................................ R EF E R EN C ES .......................................................... CHAPTER 12 LANDING GEAR SUJBSYSTEM LISTO SYM BO LS ..................................................... G EA R T Y PE S .......................................................... W H EELG EA R........................................................ G enera .......... ...... ................ ............................. Comrponent Design and Selection .................................... Tires .................. .............. .............................. W hec ls .............................................................. Shock Struts ................... ................................. B rakes ............................................................... SK ID G EA R .......................................................... G enera l ............................................................... G round-handling W he ............................ ................. Scuff P late% ....... ................................................. RETRACTABLEG EAR .................... .......................... G eneral ............................................................... A ctuatio n ............................................................. Em ergency Extension ..................................................
II-I I1-1 Il-1 11-1 I1-! l111-2 I1-2 11-2 11-2 11-4 I1-5 i-5 1l-5 11-6 11-6 11-6 .11-7
11-7 11-7 11-9 li-Il 11-12 11-12 11-12 1 1-13 1 1-13 11-13
12-I 12-1 12-1 12-1 12-3 12-3 12-4 12-5 12-8 12-8 12-8 12-8 12-9 12-9 12-9 12-9 12-9
AMI;P 706-202 TABLE OF ('ONTFNTS (('ouilnued)
Paragraph
Page
I2- 1.4 12-1.4.1 12-1.4.2 12-2 12-3 12-4 12-4.1 124 .2 12-4.3 12-4.4 124.5
R EFER EN C ES .'. ..................................
12-14
.....................
13-0 13-1
13-2 13-2.1.1
PERSONNEL ACCOMMODATIONS ...... ......................... General Vision Requirennents .................................... .1
13-2.1.2.2 13-2.1.3 13-2.1.3.1 13-2.1.3.2 13-2.1.4 13-2.2 13-2.2.1 13-2.2.2 13-2.2.3 13-2.2.4 13-2.2.5 13-2.2.6 13-2.3 13-2.3.1 13-2.3.2 13-2.3.3 13-2.3 4 13-2.3.5 13-2,3.6 13-2.3.7 13-2.4 13-2.4. I 13-2.4.2 13-2.4.3
',
12-9 12-9 12-10 12-11 12-11 12.12 12-12 12-12 12-12 12-13 12-14
CHAPTER 13 CREW STATIONS AND CARGO PROVISIONS LIST OF SYM BOLS .............. t ................................ INTRODUCTION ...................... .............................
13-2.1.2 13-2.1.2.1
)
SKIS AND BEAR PAW S ............................................. G eneral ..................................................... . ... . . Installatior .................................................. LANDING LOAD ANALYSIS AVOIDANCE O!"GROUND RESONANCE ............................. WATER-LANDING CAPABILITY ...................................... G EN ER A L ............. .............................................. PRIM E CAPABILITY ................................................. ADDITIONAL CAPABILITY ........................................ EMERGENCY FLOTATION CAPABILITY ........................... M O D ELTESTS .......................................................
13-2.5 13-2.5.1 13-2...2
i Ik,•.,% F
•II .
..
..
.
..
..
-a
•
¢...
. ..
. ..
. ..
. . . ..
.
..
.
. . ..
Controls .. Pitch C ontrols .....................................
. . ..
.
. . .
. ..
.
..
. .
.
13-I 13-1 ..
.
13-1 132 3-2
.................
Directional Control Pedals ........................................... Seats. Belts. and H arnesses ............................................. C rew Seats ........................................................... Belts and H arnesses ................... ............ ................. M ap and D ata C ases .................................................. PASSENGER COMPARTMENT ..................................... Troop and Passcngcr Seats ............................................ C o lo r ............................................... ................. U pholstering and Carpeting ........................................... Sm oking Provisions ....................................... ........... Signal Lights and A larm Bells .......................................... Acrom edical Evacuation .............................................. SURVIVAL FQUIPMENT ...................................... .. Inflight Escape and Survival Equipment ............................... Ground Escape and Ditching Provisions ............................. Em ergency L.ighting Provisions ........................................ L ife R afts ............ ................................................ Survival K its ......................................................... First Aid ................. .................................. Fire Extinguishing Svstern s and Axe .................................... ENVIRONM ENTAi CONTROL ..................................... Ventilation. Heating. and Cooling .... ................................. Windshield Defogging and Deicing Equipncnti....................... .. A coustical Environm ent ............................... ............... SIGHITS AND SIGHTING STATIONS ............................... D irect-view ing Sights ................... .............................. H elm et M ounted Sight ................................................
13-2 13-2 I -3 13-3 13-3 13-4 13-5 13-5 13-5 13-5 13-5 13-5 13-5 13-5 13-6 13-6 13-7 13-7 13-7 13-7 13-7 13-7 13-7 13-7 13-8 13-8 13-8 13-8 13-9
.TABI .01O
(11"VI'Tl'
S I (.'i1niinu.d)
Page
Parjoraph 13-2.5.3 13-2.5.4 13-3 13-3.1 13-3.1.1 13-3.1.2 13-3.1.3 13-3.1.4 13-3.1.5 13-3.1.6 13-3.2 13-3.2.1 13-3.2.2 13-3.2.2.1 13-3.2.2.2 13-3.2.3 13-3.2.4 13-3.2.5 13-3.2.6 13-3.2.7 .3-37.8
Cabin and Compartment Lighting ...................................... ................................ C ockpit Lighting ..................... Utility Lights ............................................... ..................................... Secondary Lighting ............ Panel Lightii• ........................................................ Interior Emergcncy ! ;oht . ....... .................................... Portable Inspection Lights ............................................. Troop Jum p Signal Light .............................................. Warning, Caution, and Advisory Lights ................................ |Inmtrm ent lanel lighting ........................................
13-3.2.9
Cargo Comnartment Lighting .......................................
13-4 13-4.1 13-4.1.1 13-4 .i.2 134 .1.3 13-4.2 13-4.2.1 13-4.2.2
S13-4.2.3 13-4.2.4
13-9 13-9 13-9 13-9 13-9 13-9 13-10 13-10
Indirect Sights ..... ................................................... M is~ile Sighting Stations ............................................... LIG HTING SYSTEM S .................................................. ................................... EXTERIOR LIGHTING SYSTEM A nticollision Light System ............................................. Form ation Lights ..................................................... Landing/Taxi Light ......................................... ........ Searchlight ........................................... ................ ............................................ Floodlight System Position Lights ........................................................ INTERIOR LIGHTING SYSTEM .....................................
3-10 13-10 13-10
.
....
CARGO PROVISIONS .................................................. .................... INTERNALCARGO ............................. Cargo Compartment Layout ........................................... Detail Design ......................................................... Loading A ids ......................................................... EXTERNAL CARGO ................................................. Static Loads ........................................................... ....................................... D ynam ic Loads ............... W inches and H ooks ................................................... .......... ......................... .. System Safety ............
13-10 13-10 13.10 13-10 13.10 13-10 13-10 13-I1 13-I1 13-11
13--il 13-11 13-11 13-11 13-11 13-13 13-14 13-18 13-18 13-19 13-20
t.
.--.......................... .....
CHAPTER 14 14-0 14-1 14-2 14-2.1 14 -2.1.1 14-2.1.2 14-2. 1.2.1 14-2.1.2.2
"14-2.1.2.3 14-2.1.2.4 14-2.1.2.5 14-2.1.2.6 xx
ARMOR. ARMAMENT. AND PROTE('TIVE SUIBSYSTEMS l)E);l(;N .............................. LIST O F SY M BO LS ..................... IN TRO D U CTIO N ....................................................... ..................... ARM IAM ENT SYSTEM S .......................... G U N S ................................................................. T ypes ................................................................. Location ...... ....................................................... Projectile Flight Path ................................................. Blast Effecis ........................................................ Debiis Ejection Path ................................................. External G un Jettisoning ............................................. A ccessibility ......................................................... Dynam ic Forces ......................................................
14-1 14-1 14-1 14-1 14-1 14-2 14.3 14-3 14-3 14-3 14-3 14-4
€
1iil.f f
(OF'O:NT
'ontinucd i NT( l Page
Paragraph 14-:.1.3 14 2.1.3.1 14-2.1.3 2 14-2.1.3.3 14-2.1.4 W14 2.1.5 14-2.1.6 14-2.2 14-2.2.1 14-2.2.2 14.2.2.3 14-2.2.4 14-2.2.5 14-2.2.6 14-2.2.7
Typcs of Installation; .................................................. Pod Installations ..................................................... Turret Installations ................................................... Pintle G uns .......................................................... A m m unition Storage .................................................. A m m unition Feed ..................................................... Boresighting and Harmonization ....................................... G UIDED M ISSILES ................................................... Location of Launcher Installations ..................................... Structural Clearance ................................................... Blast Protection ....................................................... A ccessibility ................ ......................................... Firing C ircuit Testing ............................... .................. Jettisoning ................................................. Effects of Aircraft M aneuvers ..........................................
14-4 14-4 14-4 14-5 14-5 14-6 14-6 14-6 14-6 14-7 14-7 14-7 14-7 14-7 14-7
14-2.2.8
Types of Installations ..................................................
14-7
14-2.2.9 14-2.2. iO 14-2.2.11
L oading .............................................................. Aerodynamic Effccts .................................................. Suspension and Retention .............................................
14-7 14-7 14-7
14-2.2.13 14-2.2.14 14-2.3 14-2.3.1 14-2.3.2 14-2.3.3
Restraining Latch ..................................................... Forced Ejection ....................................................... R O C K ETS ............................................................ Rocket Launcher Installations ......................................... Launch Tube M aterials ................................................ Launcher M ounting ...................................................
14-8 14-8 14-8 14-8 14-9 14-9
14-Z.2.i2
Launch initia.ion
"14-2.3.4 14-2.3.5 14-2. 3.6
.....
14-9
Load Requirem ents ................................................... G round Safety ........................................................
14-9 14-9
Firing Contacts ....................................................... Intervalonicter ........................................................ Launcher Fairing ............................................... ......
14-2.4.1 14-2.4.2 14-2.4.3 14-2.4.4 14-2.4.5 14-2.4.6 14-2.4.7 14-2.4.8 14-3 14-3.1 14-3.2 14-3.2.1 14-3.2.2 14-3.2.3 14-3.2.4
"14-3.2.5
14-9 1i4-i0 14-10
........................................
14-10
Safety C riteria ......................................................... Fire Interrupters ................... .................................. Contour Followers ............................................... Burst Lim iters ......................................................... C ockpit N oise ................................. ....................... Debris D isposal ................... ................................... Toxic Explosive Gas Prot,:ction ........................................ Turret M as,¢r Power Switch ........................................... PROTECTIVE SUBSYSTEMS ........................................... G EN ER A L ............................................................ DEVELOPMENT OF VULNERABILITY REDUCTION SYSTEMS.. Vulnerability Analysis ................................................ Vulnerability Reduction Checklist ................................... Vulnerability Data Presentation ....................................... Aircrew Armor Configuration Development ........................... Armor M aterial Selection ............. ................................
14-10 14-10 14-1 14-I I 14-I 1 14-I I 14-12 14-12 1412 14-1 2 14-13 14-13 14-16 14-16 14-16 14-16
SAFETY CONSIDERATIONS
'
A
4-4
...........................................
Restraining Latches
14-2.3.8 14-2.3.9 14-2.3.10
"
4-8
N umber of Rockets ....................................................
14-2.3.7
14-2.4
)
.........................................................
[
.
xxi
K t.
K
'-'.
.. . . ...
ANCP 706-PM "TABLE
OF CONTENTS (('onlinucd) Page
Paragraph 14-3.3 14-3.3.1 14-3.3.2 14-3.3.3 14-3.3.4 14-3.3.5 14-3.3.6 14-3.36.1 14-3.3.6.2 14-3.3.6.3
15-1 15-2 05.2.1 15-2.2 15-2.3 15-,.4 15 2.5 15-3
ARMOR INSTALLATION DESIGN CONSIDERATIONS ............ A irc, ew Torso A rm or ................................................. Interchangcability ..................................................... .................................. R em ovability ...................... F ,ying Qualities ....................................................... Im m obilization ........................................................ Armor Material Attachmcnt/Installation ............................... M ounting of Arm or Plate ............................................. Installation Design ............................. ..................... Bullet Splash and Spall ........... ................................... R EFER EN C ES ..........................................................
14-18 14-18 14-18 14-18 14-19 14-19 14-19 14-19 14-19 14-20 14-20
('HAPTER 15 MAINTENANCE AND GROUND SUPPORT EQIIPMENT ((;SF) INrFRFA('F IN T RO D U C T IO N ....................................................... DESIGN CONSIDERATIONS AND REQUIREMENTS ................. SA F ET Y .............................................................. A CC FSSIBILIT Y ...................................................... STANDARD IZATION ................................................ HUM AN ENG INEERING ............................................ INSPECTION, TEST, AND DIAGNOSTIC SYSTEM .................. PROPULSION SUBSYSTEM INTERFACES ............................
15-1 15-1 15-1 15-2 15-2 15-3 5-3 15-3
15-3.1
G E N E R A L ............................................................
15-3
15-3.2 15-3.3 15-3.4
INTERCHANG EA BILITY/QUICK-CHANGE ........................ CONNECTORS AND DISCONNECT POINTS ...................... INSPECTION AND TEST POINTS ....................................
15-4 15-4 15-4
OIL, FUEL, AND LUBRICATICN ................................... G RO U N D IN G ........................................................ STARTING ................................. S A R IN .......................... 15 3 . ............................. GROUND HEATERS .................. 53.8 EN G IN E W A SH ....................................................... 15-3.9 15-4 TRANSMISSIONS AND DRIVES ....................................... 15-5 ROTORS AND PROPELLERS ........................................ 15-6 FLIG HT CONTRO LS ................................................... ......................... 15-6.1 ROTATING SYSTEM S ...................... 15-6.2 NONROTATING SYSTEMS .......................................... T R IM SY ST EM S ...... . .............................................. 15-6.3 ELECTRICAL SUBSYSTEMS ......................................... 15-7 A VIO N IC SU BSYSTEM S ................................................ 15-8 15-8.1 COMMUNICATION SYSTEMS .................................... 15-8.2 NAVIG ATION SYSTEM S .. ... ...................................... 15-9 HYDRAULIC AND PNEUMATIC SUBSYSTLMS ..................... 15-9.1 HYDRAULICSUBSYSTEM ...................................... PNEUMATIC SUBSYSTEM ................................... 15-9.2 15-10 INSTRUMENTATION SUBSYSI EMS .................................. ............... 15-10.1 FLIGHT INSTRUMENTS ...................... NAVIGATION INSTRUMENTS ...................................... 15-10.2 AERIAL VEHICLE SUBSYSTI M INS1 RUMt-N]ATION ............ 15-10.3 ........... ................. A IRFRA M E STR UCTU RE ................ 15-11 LANDING GEAR SUBSYSTIM ..................................... 15-12 15-3.5 15-3.6 15-3,7
xxii
.
15-4 15-4 15-4 15-4 154 15-5 15-5 15-5 15-5 15-5 15-5 15-5 15-6 15-6 15-6 15-6 15-6 15-7 15-7 15-7 15-7 15-7 15-7 15-7 15X
L)
"
AMCP 7CO-202 TABLEl- ('ONTENINS C(ontinued) Paragraph
S"
Page
15-13
C R EW STATIO N S ...................
"15-14
ARMAMEN
AND PROTECTIVE SYSTEMS ................
15-8
16-0 16-1 16-2 16-2.1 16-2.2 16-2.2.1 16-2.2.2 16-2.2.3 16-2.2.4
STANDARD PARTI LIST O F SY M BO LS ..................................................... IN TRO D U CTIO N ........................................ r .............. FA ST FN E IRS .................... ....................................... GENERAL .................................................. THREADED FASTENERS ...................................... Screw s ............................................................... B o lts ............................. ............................. ...... N uts .. ...................... ......................................... W ashers ........................ ... .............. ..................
16-1 16-1 16-1 16-I 16-1 16-1 16-2 16-2 16-2
:
16-2.3 16-2 .3.1 16-2.3.2 16-2.3.3 16-2.3.4 16-2.3.6 16-2.3.7 -. 16-3 16-3.1 16-3.2 16-3.2.1 16-3.2.2 16-3.2.3 16-3.3 16-3.3.1 16-3.3.2 16-3.3.3 16-3 3.4 16-3.4 16-3.5 16-3.6 16-3.7 16 -3.7 .1 16-3.7.2 16-4
16-4.1 16-4.2.
r, ARMOR,
................................
('lAPTFR 1I)
16-4.2.2 16-4.2.2.1
16-4.2.2.2 16-4.2.2.3
.
NONTltREADED FASTLNERS ................................... R ivets ......................... ...................................... Pins. ...................................................... Quick-release Fasteners ............. ............................... .............. .. Turnbuckiics and Tcfinitaiil. .... Retaining Rings ........ ................................. Clam ps and G rom mets ................................................ Self-retaining Fasteners ................................................ HEARINGS............. ........................ ......I............ G E N ERA l . ................ ........................................... BA LL BEA RIN G S ..................................................... R adial Ball Bearings ................................................... A ngular Contact Bearings ............................................. Thrust Bali Bearings ... ................... ........................ ROLLER BEARINGS ................................................ C ylindrical Roller Bearings ............................................ N eedle Bearing% ............... ................ ........... ........ Spherical Roller Bearings ............................................. Tapered R oller Bearings ................ .............................. A IRFRAM E BEARIN G S .............................................. SLIDING BEARINGS ......................................... LAMINATED EL.ASTOMiRIC BI-ARINYS .......................... BEARING SEALS AND RETAINERS .............................
16-3 16 -3 K.3 16-3 16.3 b16-4 164 16-4 16-4 16-b 16-8 16-9 16-10 16-10 16-10
S eals ....... .. . . .............................. ............... Bearing R tcnrtion ..................................................... ELECTRICAL FITTINGS ...................
16 -15 16-15 16-16
16-11 16-11 16-12 16-12 16-14 16-15
G E N E R A L ....................... ... ..... .......................... CONNECTORS AND CABLE ADAPTERS ............................ Connector Selection .N... A ............................. C ircular C opnectors ................. .......... .....................
16-4.2.1
16-16 16-16 16-16
16-18
Termination Seals ............................... C able A dapters ......................................................
16-19 16-19
Connector Couplings ...............................................
16-19
16-4.2.3 16-4.2.4
Rack and Panel Connector . ...................................... Flat Conductor Cable Connector ....................................
16-4.2.5
Printed W iring Board Connector .r ................. ....................
-•--:
15-8
.
16-19 16-19
16-19
"x
xiii
Si.......
'•
I
AM CP 706-202
\EI
(%I
"l.,AHI.I. OI1 ( ()\I i '\1 SI( I ,,niiguud
! Paragraphl 16-4 .3 16-4.4 16-5 16-5.1 16-5.1.1 16-5.1.2 16-5.1.3 16-6 16.6.1 16-6.2 16-6.2.1 16-6.2.2 16-6.2.3 16-6.2.4 16-6.2.5 16-6 2.6 16-6.2.7 16-7 16-7.1 I 16.L 16-7.2.1 16-7.2 2 16-7.2.3 16-7.2.4 16-7.2.5 16-7.3 16-7.4 16-8 16-8.1 16-8.2 16-8.2.1 16-8.2.2 16-8.2.3 16-8 .2 .4 16-8.2.5 16-8.3 16-8.3.1 16-8.3.2 16-8.3.3 16"9 16-9.1 16-9.2 16-9.3 16-9.4 16-9.4.1 16-9.4.2 16-9.4.3 16-9.4.4 16-9.4.5 16-9.5
xxiv
T E R M IN Al S ... ........... ................................. ........ TERMINAl BOARDS ........................................... ELECTRICAL SW ITCHtES .............................................. ........... ........ ................... G EN I'R AI ................. ...... .............................................. T oggle S% itclics Pu.,h-button SAitches ...................................... ... . ..... ......... ....... R otaw S's itchc% ..... ............................... PIPE ANI) TUBE !ITTINc;S ......................................... GENERAl ..... ............................................. TYPES 1F: lTT NGS . . ....................................... . T apered Pipe I hreads ............ .................................... Stritght T hread Fittings ............................................... Flared T ube Fitlings .................................................. Flareless T ube Fitting% ............ .................................... Thin Wall Tube Connectors ........................................ Q uick-disconnect C oupin , ........................................... ..................... Perm anent F ittings .............................. C O N T ROL. PU LL EY S ........................ .......................... G E N E R A L ............................................................ ILLEY , -............................................................ SE.LECTIO P ulle%D iam eter ........................ ............................. Pulle) G roove ... ...... ....... ....................................... P ulle) S treng th .......................................... ............. Pulley Perform ance .................................................... ..... N onm etallic Pule) s .... ........................................ PU IL EY IN STAI LATIO N ......... .................................. PU LLEY G U A R D S ................................................... PUSH-PULL CONTROl S ANI) [ItiXiBII. SIAI"TS ................... G E N E R A L ............................................................ ......................... PUSH-PULL CONTROLS ................... ................. C ontrol T ravel ...................................... ................................ C ontrol L oads ....................... ........................... Core Configurations ................... C o n d ui. ............................................................... End F:itting, ................................................ FLEXIBLE SHAFTS ........................................... Torque C apacity ..................................................... .................... Flexible Power Shafts . .......................... Flexible Control Shafts .. ......................................... CABLES AND WIRES(STRUCTURAI ) ................................ G E N E R A L ............................................................ PREFORMED WRE STRAND AND CABL I . ....................... TYPES 01: CABLE CONSTRUCTION ............................... C A BLE SELECT IO N .................................................. C able Strength ........... ............................................ ............................................. Cable Deflection Operating Characteristics ....................................... . W ire M aterial ...................................................... C ablc(.onstruction ............. ....................... ...... ....... SAFETY WIRE AND COTTIER PINS .................................. R EFER EN C ES .........................................................
6 19 ,16-19 6-20 .I%-20 16 -20 l 20 16 -22 1622 122 16-22
.
16 22 16-23
16-23 16-24 16-25 16-25
..
.
.
.
.
16 25 l ,-25 1 6-25 ,I 2 16-211 16-26 16-2(1 16-26 .1 -26, 1(,-2(, 16-27 16-27 16-27 16,-27 I t,-2,,s 16 -29 16-28 16-28 16-21) 16-29 16-30 16-30 16-30 (1 31 16-3 1 16-31 16-3 1 16-31 1(-3?. 1(0-3I 16 32 1 -320 16 -32 16-32 16-3.
706-202
_____AAMCP
TABI.E 01 (ON'I EN• S i(',.,.ue _Paragraph
Pape (HAIA
T
llR 17
VROCl*:sSFS 17-I 17-2 17-2.1 17-2.2 17-2.2.1 17-2.2.2
IN T R O I)U C T IO N ..................... . .......... ... ................. M ET A LW O R K IN G .............. ............................. ........ G E N E R A I . . .......................................................... C A ST IN G ...... .............................. ....................... Sand C astings ............. ............ ............................. Investment Castings ......................................... ...... Permanent Mold Castin. ........................................ C entrifugal C astings ................................................... FO R G IN G ............................................................ E XT R U SIO N .................................... ..................... SHEET-METAL FORMING .......................................... M achine Form ing .................. ................................. Shop Fabrication ........ ............................................. M M............ E N IN G ........ .. ................... ................
"17-2.2.3 17-2 2 4 17-2.3 17-2.4 17-2.5 17-2.5.I 17-2.5.2 17-3
17-3.1
GENERAL................. ..................................
17-3.2 17-3.3
MACHINING OPERATIONS ...................................... ELEMENTS OF MACHINING DESIGN .............................. JOINING .......................... GEN ER A L ..........................................................
17-4
17-4.1
9.17-4.2
...
..
WELDING, BRAZING, AND SOLDERING .......................... W elding .............................................................. B razing ................... ........................................... So ldering .. ........................... ... . ........................ MECHANICAL FASTENING ......................................... R ivets ................................................................ Bolts, N uts, and W ashers .............................................. Screw s ................................................................ ADHESIVE BONDING - STRUCTURAL .......................... SWAGING AND CABLE SPI ICING .................................. IIEAT TREATM ENT ................... ............................. G EN ERAL ............................. .......... ................ HEATTFREATMtENT METALLURGY ....... .......... .. .. A nnealing ............................................. N orm alizing .. ........................................... ............ Stress Relief ........ Tem pering ............................................................ A ging ........................................ ........................ FERRO U S ALLOYS ......... ............................ ........... NONFERROUS ALLOYS ............................................. A lum inum A lloys ..................................................... C opper A lloys ......................... ............................... T itanium A iloys ................ 0 .. ................................... DESIGN ASPECTS OF HEAT TREATING ............................ WORK HARDENING .................................................. G E N ER A L ............................................................ FO R M IN G ...................... ..................................... ROLLER BURNISHING .............................................. SHOT-PEEN ING .....................................................
17-4 .2.1 17-4 .2.2 17-4 .2.3 17-4.3 17-4 .3.1 17-4-3.2 17-4 .3.3 17-4.4 17.4.5 17-5 17-5.1 17-5.2 17-5.2.1 17-5.2.2 17-5.2.3 17.5.2.4 17-5.2.5 17-5.3 17-5.4 17-5.1.1 17-5.4.2 17-5.4.3 17-5.5 17-6 17-6.1 17-6.2 17-6.3
ST7.6.4 i'
•
17-I 17-I 17-I 17-1 17.2 17-2 17. 17-2 17-2 17-2 17-3 17-3 17-3 17-4
17-4
17-5 17-5 7-6
!7-6
17-7 17-7 17-8 17-10 17-10 17-10 17-1I .7-11 17-12 17-15 17-16 17-16 17-17 17-17 17-17 17-17 17-17 17-17 17-17 17-18 17-18 17-18 17-18 17-19 17-19 17-19 17-19 17-19 17-20 xxv
TABLE OF C17N1 EN"IS (Continued) Page
Piragraph 17.7 .17-7.1 17-7.2 17-7.3 17-7.4
................................. TOOLING ........................ GENERAL ....................................................... SHOP TOOLING.................................................. ......... AIRFRAME TOOLING .................................. T EST TOOLING .................................................. REFERENCES................... ..................................
1720 17-20 17-22 17-22 17-23 17-23
APPENDIX A ]EXAMPLE OF A PRELIMINARY HEATING, COOLING, AND V'ENfILATION ANALY'1S A-1 ...... HEATING AND V~ENTILATION ANALYSIS................... A-1 A-1 DESIGN REQUIREMENTS ........................................ A-1.1 A-1 DESIGN ASSUMPTIONS .......................................... A-1.2 A-I ....... H EAT LOSSES................ ............................ A-1.3 A-1 Cockpit.......................................................... A-1.3.i A-1 ......................... Convection............................ A-1.3.1.1 A-2 Infiltration ................................................... A-l.'.I.2 A-2 Total Cockpit Heat Loss..................................... I...... A-1.3.1.3 A-2 Cabin....................... .................................... A-1.3.2 A-Z A-1.3.2.l Convection ..................................................... AA-1. 3.2.2 Infiltration ........................................................... A-', Tota! Cabini Heat Loss.................................................. A-1.3.2.3 A-2 ........ VENT ILATING AIR REQUIRED ........................... A-I1.4 A-2 A-1.4.1 Based on Number of Occupants and Minimum Ventilating Rate............ A-3 Requirement Based on Maximum Allowable Temperature Difference ... A- 1.4.2 A-3 I............................. Cockpit Requirement................. A- 1.4.2.1 A-3 Cabin Requiremcnt................................................ A-1.4.2.2 A-3 A- 1.4.2.3 Total Air Requirein~eit . .......................................... A-3 I.................................. Total Heat Requirement .......... A- 1.4.3 A-3 HEATER REQUIREMENTS................. ........................ A-1.5 A-3 ........ Heat Gained.............................................. A-1.5.1 A
V
I C~ '%
A-2.3. i
A-2.3.2
Effective A~'s..................................................
A-2.4 I
A-2 4. 1-1 A-2.4.1.2 A-2.4.1.3 A-2.4.1.4 A-2.4.1 xxvi
Iia
................... Heater Size .................................... BLOWER SIZE ................................................... Volume of Air to be Delivered ....................................... Pressure Drop . . . . . . . . . . . . . . . . . . . . . . . . . . . .. COOLING AND VENTILATING ANALYSIS .............. DESIGN REQUIREMENTS ...................... DESIGN ASSUMNiPTIONS ....................... DETERNIINOATION OF EFFECTIVE TFMiPLR,.TURIDIFFERENCLS ASSOCIATED WITH VARIOUS SURFACES OF THlE HELICOPTER .............................................. Effective Solar Tempe~ratures ........................................
A-i.5.3 A-1.6 A. 1.6.1 A-1.6.2 A-? A-2.1 A-2.2 A,-2.1
COCKPIT HEAT GAINS........................................... ionvectiori. lnfiltrat'on, and Solar Radiation........................ Convection Gains................................... ............ Infiltration Gain ...................... ............. ........... Solar Radiation Gajii ................... ........................ Total Hleat Gain Ducto Convection. Infiltrationi and Solar Radiation .. Occupants .......................................................
A~AUI~ IAI
A-4 A-4 A-4 A-4 A-4 A-4 A-4 A-4 A-4
A-5 A-5 A. 5 A-5 A-6 A-6 A-6 A-6
(
TABLE OF CONTENTS I(onfinaed) Paragraph
A-2.4.3 A-2.4.4 A-2.5 A-2.5.1
fAtC
'Electrical System .............. ........... Total Cockpit liea&. Gain .................................... AIR CONDITIONER SIZE ........................................... Conditioning of Ventilation Air ........................................
A-6 x6 A-6 A-6
A -2.5.2
Fan Size and H rit .....................................................
A -7
A-2.5.3
Tons of Refrigeration Required ....................................... -REFEREN CES .......................................................... IND•EX .................................. ..............
A-7 A -7 . J-i
.9
'22
S",,,.XXV8I
ýLIST OF ILLUSThAA 9ýNS it.,$u No,
2
Titie-ct~y
fig. 2-I Sdwc Srcue. ......... ..................... -2 Fig. 2-2 'Weight Compai ison of Matcri,0%ýor Equial 1ttii. ....... .. ......... 2-22 -Fige. 2-3 Comparative Sonic Fatigue Rnsistancý: fltZpnvcjiti~oawA bud S-4ndwichi4 '.Structures . . . . . . ... . . . . . ... . . ... .. 2-22 4Fig. 2-4 -Common Honeycomb Conf;ý urations ............................. .... 123 fig. 25 trperdecs of Balss Wood - Comprcssive Strengtl i,\ rbnsity.............. -24 t%4ig. 2-6 Properties of Balsa Wood - "L" Shear Strength vs Denwy ................ .$24 3Z tig.44 27 Typical Stabilized Comprcsaivc Strength ......... I......... 2-24 2-24 ............. ............................. Strength Typical "1" Shear ~'r. 2-8 vig. 2-9 .Typicrsl "L" Sheur Modulus........................................ (24 IFtg. 2-lb Modes fr(ailu.- of Sandwich Componimc Under Edgcwisc Loads ........... -,-.2-7 ' Fig. 3-; Submer-ged Engine Installation (Exantplc) .............................. .3-2 'wFig. 3-2 Scmicxposcd Engine lnstclladon (Example).............. ............... 3-2 Fir. 3-3 Exposed En-im: Installation (Exiampc)................. .33 KFig. 3-4 Typical Fuel Subsytcmn.................. ... .......... .............. 3-10 fig. 3-5 Typical Fuel Subsystem Wit Pressure Rcfuel~ng.......... ............... 3-12 rig. 3-6 PFerformance Cci rrczions for Duct Lovsses................. ....... .3-17 *fig.3-7 Allowable Combined lvde;ý asnd Exhaust Duct Pressure LOSSeS................ 3-17 fig. 4.1 4Helicopter Main GeiL box Weignt vs Takeoff Power ....................... 4-4 f ig. 4-2 tower Loss to Hcfat vs lnpui Povwrc -- Typical Twin -engine-driven Gearbox 4-5 ""A"............ I .......... ........ 4-7 t.ig. 4-4 Elastic Body Contact Pressurt IlIbstribution and interface Contour ........... 4$Fig.4.5 .Fri,.Iion Coeffhieint vs EI-1I) Parameters - Regions I and 11 ............. 4-8 f'ig.- 4-6 :"Angle Of Enga.;emcnt ............................................... 4-9 _l-ig. 4-7 Coefficient of Friction vs S9iting Velocity ............................... 4-10 S4-8 of &urface 'texturt and Lay on Friction and Scuffing Behavior ..... 4-10 ~ ~ ~ N~unibci of Failures vs howurs Srice Overhaul - MTBF- 500 hr............4 4-4 't' Number of Failurc:s vs HWoes &nuriv Operation- MTBF 5000 hr........... 4-14 m.-I 4rotm~aily of Survival vs L/rF,)Ratio...................... ............ 4-15 r'Cig j i'ý ~ pli~ivs Hertz Stress...........................................41 F.i 4 -Ii 14cibull Plot - Spalling Life vsGear Population Rank .................... ...... 4-"I ig44 'T pica 1 aiVPRotnýr Gearbox - Vulnerable ...................... 4A-19 ~f%4 '~~iUoto~earo-%X n zxo ....-. ...... 1........................4-% jW IC Zvý 0 ote,; Gcarbois - 12.7 Mmn Proof.............................. 11-21 .0,1% Typicwil Spects 1Powe.rFuncticn ......................................... 4-25 4S !2ý Li~~ee~ feCultvc,; ............................. ....... 42 I 4j4ig419 Shaft Horspowei ~cr His-Lograms................... ......... 4c$I~4-20 Orioph cRlationship - Fwilnte- Modes - Load v., 'eoi hy ............. 4-35 -A-f iraphic- Rel~ationship -- Failure Modes - Load vs Tjooth Siie........ 4-35 'Fiag. 4-22 . S5ingki VYoat!) Pusorf Gear Fatigue Test Results ........... I......I........ 4-41 f4. *-23 z vs SýWini: Velocity - Synchroniizd ani Ui~ytchreiiae Dics 442 -Tig4ý--_, C4i L -,,týiAtV>'wabke *i Suibsurfacxý Shear........ ........... ....... 4-45 4-46 -* Unsynchror.z. ................................ s V50tr.:b'cnt 4?i
.
34
.............
:C.Tect~
,~
.
.
:
-4-26;~
j
-0ý1
2 4-
Cr~~ ear--Inrc~Ring Fit vw peratinTime ................ ........ c s ater Ring -ILinc~r Fit ReC$,4tion...... ............... -.:nGtigy 4 With ittcer Ring Expansion .................. -.-
7zrn al lieoad-
OF vs. Dl?.....................
Sp~t!nul on lh'Ttuhant Bearing fotccs................... Clearmnat LLv - lasic.anti -syj 4
..
1
4-49 4
45
-56
4
~-A
LIST 01- 1 I.V34TA~ll .o 0 0- Caef~iwwd) Fig. No,
*
Fig. 4-33 F-ig. 4-34 Fiwg 4135 Fig. 4-36~ Fig. 4-37 Fig- 4-38 Fig. 4-39 Fig. 4-40 Fig. 4-41 Fig. 4-42 Fig. 4-43 Fig, 4-44 Fig. 5-1 Fig. 5-? Fig. 5-3 Fig. 5-4 Fig. 5.5 5-6 Fig. 5-7
IFig.
RP
iý 5-8
*
Involute Spline Data 1, vs;N ... ....... ............ 4-60 Radial Mode Resonance 17ý_owiic vN Cear t'ooth Meshing Sipted ...... 41 Typical Spiral Damper Ring App~k~aions ............. .......... ........ 4-70A Relative Shaft Speed vs Pantive Vib~ration Amplitude......................4-74 Typical Bcaring Flangc~r As-_cmbl.V - Subci itical Shaf! Assembly .............. 4-75 Flc'-ible Diphragm Coupirr,... .................... ...........I... 4-77 Bosskcr Coupling ....................................... ............ 4-18 Elastom,:ric Coupling ................. ............................ 4-78 Hooke's Joint (Universal) ......................... ................. 4-9 Gear Coupling ................................................ ...... 4 Breakraway Sliding Foice vs Misalignment for Various Spline Devices ..... 4-79 Oil System Seheniawc.............I........... ............ ......... 49 Vectot Diagram of Swirl in 1-over ...................................... 5-1 Control Momyent for Basic tRotor Types. ..................... .......... 54 Articulated Rotor Schematic...................................... 5-9 Coincident Flap and Lag Hinge Rotor................................... 5-10 Girnb3led Rotor Schematic................ .......................... 5-if) Tcctering Rotot Schematic .............. ... .......................... 5-Il Teetet ing Rotor..................................... ................ 5-1; i-ingcicss Rotor Schcunaiti:.......................................... XH-51 Rotor System .................................................
5-13
Fig.5-11 Fig.5-12 Fif$. 5-13 Fig. 5-14 Fig. 5-15 Fig. 5-16 Fig. 5-17
Mechnismof Pitch-lag Instability.......... ......................... Pitch-nlap Coupling of Rotors ...................................... Mechanism of Flap-lag Instability ...................................... Typical Plots of Rotor Natural Frequency vs"OperatinF Speed ............... Single-degree-of-freedom Cole-man Plot............... ................ Two-degree-of-freedom Coleman Plot ............................... ... Two-degree-of-fteedvim Coleman Plot Showing Se~tisfaction of Minimum Frequency Criteria fur Two-blade IlingEcless Rotor...................... Gus* Load FactorComputec! for the UII-IB Helicopter Using I.ineai-Theory Gust-alleviation Factor (M IL-S-8698)................................. Rotor Limits as a Function of Advance Ratio........................... Results of a Load Gust Study Compared With Military Speciricauion Requiruments.......................................... Arliculatcd Rotor (Boeing Model 107) ............ ........ ............. CH--46 and CH-47 Tension-Torsion Strap Assemblies ...................... Torsionally Stiff and Flexible Wire-'vound Tic-bar Assemblies............... Elastomeric Bearings ................................................. Hydraulic Lag Damper............................................. CH-46 Power Blade Folding Mechanism................................. Typical Helicopter Rotor Blade Airfoils.................................. Track WiVth Varying rpm (Zero Collective Pitch) ........................... Tr-ack With Varying Collective Pitc!h (Constant Rotor rpm) ................. Alternating Stress Superimposed on Steady Stress ......................... Alternating Stress vs Cycles at Various Steady Stress Levels (Crv!ss Plotted from Fig. 3-12, MIL-HDBK- 17 for Notched Specimens of 181 Glass Fabric With MIL-R-7575 Polyester Resin) ................................. Propeller Flow Field for- Compound Helicopters........................... Comparison of P-order Excitations ......................................
51 i 5-16 5-16
Fig. 5-22 F;g. 5-23 Fig. 5-24 Fig. 5-25 Fig. 5-26 Fig. 5-27 Fig. 5-28 Fig. 5-29 Fig. 5-30 Fig. 5-31 Fig. 5-32 Fig. 5-33 Fig. 5-34 xxx
Title
Fig. 5-9
Fig. 5-18 Fig. 5-19 Fig. 5-20 Fig. 5-21
.
~
5-20
5-21 5-21 5-22 5-24 52 5-25
5-25 5-28 5-32 5-33 5-33 53 5-37 5-40 5-49 5-49 5-54 5-55 5-58 5-58
4.-
P 706- 20
4.,
LIST OF ILLUSTR ATIONS (Confinued)
Propeller IP Loads from Nonaxial Inflow.............................. 1IP Excitation Diagram for Typical STOL Aircraft......................... 1IP Excitation Diagram for Helicopter With Pusher Propeller................ Propeller Critical Speed Diagram ...................................... Propeller Vibration Modes ........................................... Stall Flutter Design Chart ............................................ Airfoil Characteristics and Stall Flutter ................................. Propeller Control System Schemnatic.................................... ............................ Simplified Propulsion System Block Diagramn Linearized Propellev Control Block Diagrai'n............................. Typical Blade Cross Sections.......................................... Typical Spar-shell Blade.............................................. Blade Materials and Weight Reduction ............................... Fatigue Strength Diffcrence Between Specimen and Full-scale Tests...... ..... Typical Stress Summary Curves ....................................... Geometric Da3ta.................................................... Fin Separation Distance/Rotor Radius ................................. Sideward Flight Velocity ............................................. Tail Rotor Performance, Four Blades................................... Typical Variation in Tail Rotor Noise Level.............................. Cutimpriisati-on" Ilgtie PD,-h n- Moment With Conine Angle and Blade CG....................................................
5-58 5-59 5-59 5-60 5-61 5-64 5-64 5-67 5-69 5-70 5-71 5-72 57 5-76 5-77 5-78 5-79 5-80 5-80 5-81
Fig. 6-I Fig. 6-2
Typical Control Function Scheduling for a Tilt-rotor Aircraft................. ....................... Characteristic Root Plot............ ............
6-3 6-5
Fig. 6-3 Fig. 6-4 6-5 Fig. 6-6 Fig. 6-7 Fig. 6-8 Fig. 6-9 Fig. 6-10 Fig. 6-11I F ig. 6- 12 Fig. 6-13 Fig. 6-14 Fig. 7-1 Fig. 7-2 Fig. 7-3 Fig. 7-4 Fig. 7-5 Fig. 7-6 Fig. 7-7
Allowable Pitch Czintrol System Residual Oscillations ..................... Control Mixing Schematic ............................................ Mechanical Mixing Assembly......................................... Powered Actuators (Tandem Helicopter)............ .................... ....... Artificial Feel and Trim Schematic.............................. Rotating Controls .................................................. Typical Pitch '-ink Rod End .......................................... Centrifugal Force Deflections......................................... Pitch Link Adjustment Provisions ..................................... Relative Pitch Link Rod End Position................................... Instrumented Pitch Link ............................................. ..................................... 1Instrumentcd Drive Scissors .... Typical DC Power Distribution System ................................. Typical AC Power Distribution System ................................. Example Load Analysis AC Left 1-and Main............................. Examiple AC Load Analysis Format ................................... ........... Typical Automation Flow Chart ........................... Typical AC Generator With Oil-lubricated Bearings....................... _............ DC Starter/Generator .................................
6-8 6-IS 6-15 6-16 6-17 6-19 6-20 6-20 6-20 6-21 6-22 6-22 7-3 7-4 7-5 7-6 7-7 7-8 7-10
Fig. 5-35
~Fig. 5-36 ;
*Fig. *Fig.
Fig. 5-37 Fig. 5-38 Fig. 5-39 Fig. 5-40 Fig. 5-41 Fig. 5-42 Fg5-43 Fig. 5-44 Fig. 5-45 5-46 5-47 Fig. 5-48 Fig. 5-49 Fig. 5-50 Fig. 5-51 Fig. 5-52 Fig. 5-53 Fig. 5-54 Fill. 5-55
)Aft
*Fig.
-AFig. )
5-002
Clast-cooled DC Generator ...........................................
7-12
Fig. 7-9 Fig. 7-10 Fig. 7-1l Fig. 7-12
DC Starter Motor With Solenoid-operatcd Switch ......................... Prolotype Cartridge-boosted Electrical Starter Systemn..................... Sample Set of! tilization Loads ....................................... Gases Emitted from Nickel-Cadmium Sintered Plate Cell During Overcharge
7-12. 7-13 7-18
Fig. 7-13
PermiisihleClamp Deformation
7-8
*
Page
Title
Fig. No.
at
.7-19
70'-75*F.........................................................................
............
.........................
7-2 5
a.
AMCP 706-201 LIST OF ILI-LJSTRATlONS8Conlinued)
Fig. No.
Title
Page
Fig. 7-14 Fig. 7.15 Fig. 7-16 Fig. 8-1 Fig. 8-2 fig. 8-3 Fig. 9-I Fig. 9-2 Fig. 9-3 Fig. 9-4 Fig. 9-5 Fig. 9-6 Fig. 9-7 Fig. 9-8 Fig. 9-9 Fig. 9- 10 Fig. 9-l11 Fig. 9-12 Fig. 9-13
Terminal Strip Installation............................................. Typical Connection to Grounding Pad................................... Typical Lightning Electrical Circuit Entry Points .......................... Block Diagram of Classical Communication System ................. ...... Typical Intercommunication Selector Box ................................ Typical Cnimunication Antenna Layout.................... -.... I...... Central Hydraulic System ........... .................... ............. Dual System Hydraulic-powcred Flight Control Actuators .................. Dual-powcrcd Stability Augmentation Systcm............................. Dual-po%-:rcd Stick Boost I ydraulic System.............................. Hydraulic Starting; Energy-limited System ................................ Hydrauix Starting. Power-iirnited System............. ................... APU Starting System................................................. Cargo Door and Ram~p System ......................................... Cargo and Personnel Hoist (Constant Prcssurt) System ..................... Rotor Brake System .................................................. Wheel Brake System .................................................. Combined Spool Switching Vakve ....................................... Pressuie Check Valves Plus Power Return Switching........................
7-26 7-27 7-31 8-4 8-4 8-13 9-I 9-2 9-2 9-2 9-2 9-3 9-3 9-3 9-4 9-4 9-5 9-8 9-8
Fig. 9-14
Pressure Check Valves Plus Inline Return Relief Valve......................
9-9
Fia 9! 5
Fig. 9-16 Fig. 9-17 Fig. 9- 18 Fig. 9-19 Fig. 9-20 Fig. 9-21 Fig. 9-22 Fig. 9-23 Fin. 9-24 Fig. 9-25 Fig. 9-26 Fig. 9-27 Fig. 9-28 Fig. 9-29 Fig. 9-30 Fig. 9-31 Fig- 9-32 Fig. 9-33 Fig. 9-34 Fig. 9-35 Fig. 9-36 Fig. 9-37 Fig. 9-38 Fig. 9-39 Fig. 9-40 Fig. 9-41 Fig. 9-42 Fig. 9-43 Fig. 9-44 xxxil
Irinein Mehnicafy LockeC-out
-cifVie................
Cam-operated Poppet Switching Valvec....... ................... ....... Switching Valve...................................................... Hydraulic System Ground Fill Provisions ................................ Rosar, Boss Fitting ................................................... Use of an Articulating Link ................... ................ ........ Use of Protective Cover on H-oses ....................................... Typical Mission Requirement Profile .................................. Examples of Parallel and Series Control Modes............................ Schematic of Jet Pipe Electro-hydraulic Control Valve ..... ................ Schematic of Flapper Electro-hydraulic Control Valve...................... Typical Master Control Valve.......................................... Anticavitation Approaches ............................................ Feedback Techniques ....................... ......................... Hydraulic Pump Flow vs Pressure Characteristics .......................... Hydraulic Pump Soft Cutoff Characteristics .............................. Hyaraulic Pump Case Drain Flo%%Characteristics ......................... Suction Line Length - Reservoir Pressure Characteristics............ ...... Hydraulic Pulsation Suppressor ........................................ Filler Element Dirt-holding Characteristics ............................... Filter Element Performrince ........................................... Hydraulic Valve "Trail" Configurations ................................. Hydraulic Valve Configurations........................ ................ Direct-operated Valve ......................................... ....... Pilot *operated Valve ................ ................................. Valve Operating Time................................................. Solenoid-operated Valve Incorporating Rcl urn Pressure Sensing .............. Power and Spring Main Section Valve Return to Neutral.................... Typical Separate Servo Actuator ........................................ Dual Seals With Return Vent ............................ ..............
-
9-9 9-10 9-1l 9-1l 9-12 9-12 9-12 9-14 9-IS 9-IS 9-15 9-17 9-17 9-18 9-18 9-18 9.19 9-19 9-21 9-21 9-22 9-22 9.23 9-23 9-23 9-24 9125 9-26 9-27
cP
LIST OF ILLUSTRATIONS (Continued) Fig. No.
Title
Page
Fig. 13-8 Fig. 13-9
Methods of Raising the Suspension Point .. Helicopter Load Dynamics Schematic .................. Study Input Variables ................................................. Equivalent Steady Load for Combination of Steady and Vibrator), Loeds (Nonreversing) .................................................... Equivalent Steady Load for Combination of Steady and Vibratory Loads (Reversing) .................................. Mounting of Duplexed Bell Bearings..................................... Spherical AicatB aig .........................
13-16 13-18
16-6
Common.....f onetos........................
16-18
*Fig.
14I Fig. 16-1 Fig. 16-2
Fig. 16-3 Fg 164
Fig. 16-5 Fig. Fig. Fig. Fig.
Fig. Fig. Fig. Fig. Fig. Fia.
16-6 16-7 16-8 16-9 16-10 16-11 16-12 17-1 17-2 17-3
Fig. 17-4 Fig. 17-5
~i.176 * * *
Fig. Fig. Fig. F.g. Fig. Fig. Fig.
17-7 17-8 17-9 17-10 li-Il 17-12 17-13
Fig. 117-14 15 Ig
xxxiv
(
014-17 C
16-7 16-9 16-13
Tapered Pipe Thread Fittings........................................... 16-22 Straight Thread Fittings ............................................... 16-23 Flared lube Fittings.................................................. 16-24 Flareless Tubc Fittings ................................................ 16-24 Cable Alignment and Pulley Guai d Location.............................. 16-27 Pus~i.pull Cablus and End Fittings..................................... 16-29 Right-handed Thread Application of Safety Wire......................... 16-33 Standard Bend Radii PrActice -Minimum Bend Radii.......I...... I........ 17-5 Weld Contour and Stress Concentration ....................... 17-7 Welding Symbols.......................................... 17-8 Reprcsentative Buit Juivinb................................................ 17-8 Representative Corner Joints........................................... 17-9 Representative Tee Joints..............................................19 Rivet Spacirg ........................................................ 17-10 Types of Loading for Bonded Joints..................................... 17-12 Lap Shear Joint Deflection Under Load .................................. 17-13 Typical Rotor Blade Design - Alternate I................................ 17-13 Typical Rotor Blade Design - Atlernate 2................................ 17-13 Hioneycomb Sandwich Structure ........................................ 17-14 Addition of Doubler-' to Honeycomb Structure............................ 17-14 Balance Bar Design ............................. ..................... 17-15
as........eWngRr
Fig. 17-16
I...................
*2
.r
--
D....................
...........
Cable Splicing..................................
... ..
..
.. ..
.....................
176
17-16
-
". 't
LIST OF TABLES Table No.
Title
Page
TABLE T"ABLE YABLE TABLE TABLE
2-I 2-2 2-3 2-4 2-5
Mechanical Propcrtii.s of 18 Ni Maraging Steels ............................ Comparative Mechanical Properties for Selected Nonferrous Alloys ......... Grouping of Metals and Alloys (MIL-Si D-889) ............................ Position of Metals in the Galvanic Series ................................... Process Comparison Guide for GRI' Laminates ............................
2-4 2-4 2-7 28 2-14
TABL E 2-6
General Properties Obtainable in Some Glass Reinforced Plastics ............
2-15
TABLE 2-7 TABLE 2-8
TA BLE 4-2 TABLE 4-3 TABLE 4-4 TA BLE 4-5 TABLE 4-6 TAB! F 5-! TAB[E 5-2 TABLE 5-3
Common Resin Rcinforcemen't Combinations of Thermoset Laminates ...... 2-16 Typical Values of Physical and Mechanical Characteristics of "Reinforcement Fibers ................................................... 2-18 Nominal Composition of Glass Reinforcements ............................ 2-19 Typical Unidirectional Composite Properties Based on Commercial Prepregs 2-20 Properties of Rigid Foams ................................................. 2-25 Common Adhesives in Current Use ....................................... 2-26 Shear Bond Strengths of Adhesives ..................................... .. 2-26 Useful Temperature Range and Strength Properties of Structural Adhesives .. 2.2k Armor Material Design Data and Physical Charactetistics .................. 2-28 Fabrication Data for Lightweight Armor Materials ........................ 2-2') Typical Properties of Commonly Used Structural Adhesives ................ 2-32 Helicopter Lubricants and Hydraulic Fluids ................................ 2-3) APU Types for Main Engine Starting Environmental Control, and E lectrical Supply .................. .................................... 3-16 APU Reliallity ..................................................... 3-20 U S Army Heliconters - Transnmission ,ind nrive-Sytem. On.y 4-17 Maintenance Workload .......................................... External N oise Level ......................... .......... ..... ........... 4-18 He!icopter Drive Subsystems - Single Main Rotor. ....................... 4-24 .. 4-33 Life Modification Factors - Surface Durability ......................... Shear Stress vs Depth ............... ...................... ............. 4-45 Helicopter Transmission Case Materials and Application Data .............. 4-68 The Relative Effects of Various Parameters on Gust Response ............... 5-26 Example of Nominal Weight and CG Locations ........................... . 5-48 Rotor Blade Balance (Sample) .......................................... .. 5-48
TABLE 5-4 TABLE 5-5
Comparison of Material Properties ................................... Aerodynamic Characteristics of Several Airfoil Sections Suitable for
TABL -5-6 TABILE 6-1 "TABLE7-1 TABLE 7-2 TABLE 7-3 TABLE Il-I TABLE 11-2 TABLE 12-1 T A BLE 13-1 TABLE 13-2 TA BL E 14-1 TABLE 14-2 T A B LE 14-3 TABLE 16-1 TABLE 16-2 TABLE 16-3
Summary of Tail Rotor Excitation Sources ................................ Maximum Amplitudes of Limit-Cycle Oscillations .......................... Outputs of Converters Relative to Continuous-Current G-ncrator ........... Typical Characteristics of 24V. 34 All Battery Systems ................. ... Alternative Charging Methods ....................................... Cost Impact. Airframe Detail Design ..................................... M aterial Selection - Airframe Design .................................... Load Factors for Helicopter Tire.s ......................................... C oefficients of Friction .......................................... ......... Standard Cargo T iedessn Devices .......... ....... ....................... Typical H elicopter G uns .................................................. Vulnerability Damage Criteria Data Sumnmar% ............................ Vulnerability T able ....................................................... Life I actors for Antifriction Bcaring M aterizis ............................. Cost vs Tolerance Class for Antifriction Bcari.igs .......................... Standards for Airfrname Control Annular Ball bearing. .....................
"TABLE 2-9 TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE 'ABLE
,'
2-10 2-11 2-12 2-13 2-14 2.15 2-16 2-17 2-18 3-1
TABLE 3-2 TABLE 4-1
T ll
__\
II
RiB.tU d esic.........
...................................
2
5-51
. .........
5-82 6-8 7-15 7-16 7.17 . -. , 11-3
12-4 13-14 13-15 4-2 14-14 14-14 16-6 lo-7 16-12 \•
mm
-9.•
'••
•,.,•,
•
.,.:.:.••.
..
S.11
AMOP MG-202 I.T01:"1AI ABIS ("outinutcd
Table No. TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE TABLE
16-4 16-5 16-6 16-7 16-8 16-9 16-10 16-I 16-12 17-1 17-2 17-3 17-4
TABLE 17-5 TABLE 17-6
*
_
(.p
Titlc
Pagc
Standards for Airframe Control Rod End Bcaring.i ......................... Standards for Sphe'ical Roller Airframe Bearings .......................... Properties of Sliding Bearing Materials for Airframe Use .................... Specifications and Standards for Self-Lubricating Slide Bearings ............ Militery Specifications and Standards for Connectors for Aircraft ........... Other Military Spccificationi and Standards for Connectors ................ Military Specifications and Standards for Crimp-Style Terminals ............ Military Specifications and Standards for Switches ......................... Military Specifications for Cables .................................... Bentd Characteristics of Selected M etals .................................... Unit Horsepower Values for Represerntative Metals ....................... Representative Surface Finishes Obtained in Machining Operations ......... Values To Be Added to or Subtracted from Base Dimension for "Holesand Shafts To Calculate Tolerance ................................ Representative Heat Treat Temperatures .................................. The Effect of Shot-peening on the Fatigue Properties of Selected Samples
16-12 16-12 16-13 16-14 16-17 !(-18 16-20( 16-21 16-31 17-4 17-6 1'7-6
, -1
17-7 17-i 17-21
I
S
-,-.
\\~
!. -
1 ~-~-
AMCP 706,201
The IIhli,•p:er Lngineering llandbook form. ai part-of the 'ingincering Design Handbook Series shich presents engineering data for the design and construction of A -y equipment. This volume. AMCP 70M-202. Delail De.%ign. is Part Two of a three-part Enginering Design Handbook titled Helicopter Engineering Along with AMCP 706-201. Preliminary Design. and AMCP 706-203, Qualification A.tsurance, this part is intended to set forth explicit design standards for Army helicopters, to establish qualification requirements. and tu provide technical guidance to helicopter designers, both in the industry and within the Army.
a
This volume, AMCP 706-202, deals with the evolution of th', vehicle from an approved preliminary design configuration. As a result of this phase of the developmcnt. the design is describcd in sufficient detail to permit construction and qualifica"tion of the helicopter as being in compliance with all applicable requirements, inchiding hce approved system specification. Design requirements for all vehicle subsystems are included. The volume concists of 17 chapters and the organization is discussed in Chapter 1. the iutroduction to the volume. AMCP 706-201 deals with the preliminary design of a helicopter. The characteristics of the vehicle and of the subsystems that must be considered arc described.
ir
and possit '- solutions at ' suggested. The documentation necessary to describe the
J
"preliminarydesign
in sufficient detail to p,-rniit evaluation and approval by the procuring activity also is described.
A
The (hird volume of the handbook, AMCP 706-203, defines the rcquirements Ifor airworthiness qualification of the helicopter and for demonstration of contract cornpliance. The test procedures used by the Army in the performance of those additional tests required by the Airworthiness Qualification Program to bt performed by the Army also arc described.
xxxvii
F.
AMCP 706-202
PREFA(I: This volume, AMCP 706-202, Detail Design. is the %econd section of a three-part -ngitieering handbook, Heficopier Engineering. in the Engineering D'%sign Hand"book series. It was prepared by Forge Aerospace. Inc.. WAshington. D.C., under subcontract to the Engineering Handbook Office, Duke University, Durham. NC. The Engineering Design Handbooks fall into two basic c-'tegorics. those approved fot release and sale, and those classified for security reasons. The US Army Materiel Commano policy is to release these Engineering Design Handbooks in accordance with current DOD Directive 7230.7. dated 18 September 1973. All unclassified Handbooks can be obtained from the National Technical Information Service (NTIS). Prozedures for acquiring these Handbooks follow: a. All Department of Army activities having need for the Handbooks must submit their request on an official requisition form (DA Form 17, dated Jan 70) directly to: Commander Lettcrkenny Army Depot ATTN: DRXLE-ATD Chambcrzburg. PA 17201 (Requests for classified documents must b,. submitted. mith appropriate "Need to Know" justification, to Letterkenny Army Depot.) DA activities will not requisition Handbooks for further free distribution. b. All other requestors. DOD, Navy. Air Force. Marine Corps. nonmilitary Government agencies, contractors, private industry. individuals, universities, and others must purchase these Handbooks from:
.
National Technical Information Service Department of Commerce Springfield, VA 22151 Classified documents mey be released on a "'Need to Know" basis verified by an official Department of Army representative and processed from Defense Documentation Center (DDC), ATTN: DDC-TSR, Cameron Station, Alexandria, VA 22314. Users of the handbook are encouraged to contact USAAVSCOM, St. Louis, MO, System Development and Qualification Division, with their recommendations and comments concerning the handbook. Comments should be specific and include recommended text ¢hangcs and supporting rationale. DA Form 2028, Recommended Changes to Publications (available through normal publications supply channels) may be used for this purpose. A copy of the comments should be sent to: Commander US Arny Mateviel Development and Readin•es Command Alexandria, VA 22333 Revisions to the handbook will be made on an as-required basis and will be distrbutcd on a normal basis through the Letterkenny Army Depot.
t
AMCP 706-202 (CHAPTER I
INTRODUCTION AMCP 706-202, Engineering Design Handbook, Helicopter Enginre.-ing. Parr 7wo. Detail Design, is the second part of a th:¢¢-volume hlicopter engineering design h~ndbook. The preliminary design (covered in AMCP 706-201) is *vclopcd during the proposal phase. at which time all subsystems must be defined in sufficient detail to determine aircraft configuration, weight, and pcrformance. The detail design involves a reexamination of all subsystems iri order to define cach clement thoroughly with the aims of optimizing the aircraft with regard to mission capability as well as cost considerations. Detailed subsystem specification requirements are the basis for in-depth analysis and evaluation of subsystem charactcerstics and interfaces. Based upon complete system descriptions and layouts, performance, weight, end cost trade-offs arc finalized. Periodic reviews of the design are conducted to evaluate mairtainability, reliability, safety, producibilhty. and .oriforniancc with spc.i.fication requiremeents. Development testing may be required to permit evaluation of alternate 5olutions to design problems or to obtain adequate information for trade-off investigations. Appropriate consideration of human engineering factors often requires evaluation of informal mock-ups. SWeight control is an important element of the detail design phase. Subsystem weight budgets, prepared on the basis of the preliminary design group wCight hreakdown, arC aasiýzsed at the initiation of
the detail design phase. The continui-ig evaluation of compliance with the budget as an essential part of the manzgement of the project and the awsurancc of cornpliance with weight guarantees of the helicopter detail specification are described in conjunction with the discussion of the Weight Engineek'ing function in ANICP 706-201. The requirements and procedures for airworthiness qualification and proof of contract compliance for a new model helicopter for the US Army are defined and discussed in AMCP 706-203. which ikthe third volume in this handbook serits Qualification is not time-phwed. but is a continuing part of the acquisition program. A number of qualification requirements are integral parts of the detail design effort. Desipr. reviews by the procuring activity are re(quired during the definition of subsystem configuraons as well as during the final design of assemblies --.
and tinstallations. Evrluation of a full-scale mock-up of the complete helicopter is a major part of the design review process. The requirements for this review are described in dctail in AMCP 706-203; but the construction and inspection of the mock-up must be completed at the earliest piactical point of the detail design phase to permit the contractor to complete the desibn and manufacture of helicopters for test and for operational deployment, with reasonable, assurance tOat the configuraion is responsive to the mission ecquircments. Also completed during the detail design ohase are a variety of analyses necessary to substantiate the comp!ianci of the physical, mechanical, and dynamic characteristics of subsystems and their key cornponents with applicable design and performance requirements, including structural integrity. The analysis required during the design, development, and qualification of a given model helicopter are those specified by the applicable Contract Data Requireicis List ('Dr , L-). This volume reviews the functions pcrformre by the major helicopter subsystems and outlines the requirements app!icaolr to the design and installation of eaich one. Principal documentation of the detail design phase is the final drawings of the helicopter in sufficient detail for procurement, fabrication. assembly, and installation. This volume, therefore, aiso iincludes discussion of materials and processes pertinent to the ),-nnufacture of bclicopter components. This volume is intended to pruvide designers. engmnrr,
relaltvelv
rew to the
helicorptev tech-
nology, and program managers a general design guide covering all of the helicopter detail design specialties. however, it is not intended as a source of detailed design procedures for use by the experlcnced design engineer in his specially. Throughout this volume the mandatory design reouiremcnts have been identified with the contractual language which makes use of the word ".shall". To assist in the use of the handbook in the planning or conduct of a helicopter development program, the word "shall" has been italicized in the state.ment of each such requirement. Since: the mission requirements for individual helicopters result in variations between subsystem configurations and performance requirements. the procuring activity will specify in its Request for Proposal (RFP) the extent to which the design requirements of this handbook are applicable to the acquisition of a given helicopter. I-I
.. • ,
AMCP 7_06-202
CHAPTER 2
MATERIALS
)
2-1 INTRODUCTION
factors, standard mill products, and cost data. tl,.IL-
This chapter addresses the properties of the various materials used in the construction of helicopters, These materials include ferrous and nonferrous metal. nonmetallic materials, composite structures, adhesives and seplants, paints and finishes, lubri., cants, greases, and hydraulic fluids. Among the ferrous me:als are carbon steel, stainless steel, and al:oy steels. The nonferrous mttals include aluminum, magnesium and titanium alloys, beryllium, copper, brass, and bronze. Thermoplastic and thermosetting plastics, elastomers, woods, fabric, and fluoroplastics are reviewed. Composite structures, including filament laminates, fabric laminates, and filament wound, and honeycomb and sandwich construction ake discussed aý are the adhesives used for bonding of primary structure, honeycomb and compo3ites, fahris, rbhhr, e'astomers, g1.s; and rlastics. Sealhng compounds, such as putty and pastc;, also are de.ai!ed. A discussion of paints and coatings, special finishes, plating, and tapes is included, as well as a review of th,. most commonly used lubricants and their applications. The designer will find that a good working relationship with vendors wii! help him to keep abreast of new rtiatvrials and processes with possib!t applications to helicopters. New materials are being iiitroduced continually, and new processes alter the cost and performance relationships among older
and provides additional detail design data. For technical data and information pertaining to wrought iron, carbon steels, and low-alloy steels, an exccllent source is N4IL-IIDBK-723, which covers some of the more practical aspects of -ecta! forming and joining. Finally, for design data and metallurgical details, the designer should consult the various American Society of Metals (ASM) Handbooks. Because of weight considerations, it is desirable to restrict the heavier ferrous metals to those applications where very high strength, a high modulus of rigidity, high resistance to fatigue, and high modulus of elasticity are reqaired. The more expensive high-performance steels often are more economical in terms of weight, cost, and fatrication processes than are the lower-cost ferrous products. Anplications for these materials include high-sirt-ss parts such as rotor drive shafts, masts, hubs. vertical hinges, flapping hinges, tie cables, ti'bular frames, an6 control cams, keys, gears, and hydraulic cylinde-.s.
, -.
.
HDBK-5 is a source of mechanical properties data
..
conform to one or more Government specifications, and many manufacturers take step, to keep their products on the qualified-products lists, where such are requiret. Propei regazd for and awareness of these concerns in design callouts will simplify procurernent, fabrication, and qualification of hard•arc.
2-2 METALS FERROUS METALS
2-2.1
R"
2-2.1.1 General Thik dis•:ussion provides a brief review of ferrous metalk and their application to the construction of helicopters, as well as of some of the parametcrs go)vernling the choice of a particular ferrous metal for specifi as a AP•'','•, moreconpreheisivc discussion, as well as detail \,dcsign data, will be found in Chapter 9, AMCP 70%hich describes such items as mater'al selection r
2-2.1.2 Carbon Steels The carbon steels are a broad group of iron-base alloys having small amounts of carbon as their principal alloying element. Commonly, the carbon content falls between 0.03 and 1.2%. The American Iron and Steel Institute (AISI) codt usually is used for designating s•ate ,. Tins ...ystm i.i..i Uup.. a Or five-digit number to designate each alloy, with the first digits referring to the alloy and the last two digits giving the carbon content in points of carbon. where one point is equal to 0.01%. Thus, 1045 steel is one of a series of nonsulphurized carbon steels and has 0.45% carbon. Other carbon steel s.ies are the II XX series, which are resulpherized; the BIIlXX series, which are acid Bessemer resulphurized, and the 12XX series, which are rephosphorizcd. Low-carbon stee's range from 0.05 to 0.30% carbon, mediumcarbon steels range from 0.30 to 0.60% carbon, and high-carbon steels range from 0.60 to 0.95% carbon. The machinability of low-carbon steels is poor. They tend to drag and smear and to build up on the cutting edges of tzols, generating considerable heat and decreasing cutting cffiricncy. Medium-carbon steels machine .better, although the cutting pressures are higher. High-.arbon steels are too hard for good machining, but they are used where fine finish and dimensional accuracy are required. Hot- and coldfuui--
accrac
27I
-:
-'
AMCP 706-202 rolied stels machinc better than do annealed steels, and the macbining properties of the low-carbon steels arc improved by 1he addition of sulphur. phosphorus. or lead. The low-carbon steels have excellent forming properties, and can be worked readily by any of the normal shapin3 processes. Their ready formability is due to the fht that there is less carbon to interfere lans th sam tokn, heidr-. withthe ofslip witulty plao increafsesp. witheinmereasing carbon df ok i a wtough ctn content. Plain carbon steel is the most readily welded of all materials. Low-carbon (0.15%) steel presents tke least
rosion resis,.anc is about the same. hih I lie AISI designation system is used for alloy steels also. This is illustrated by 4130 steel, -hich Is an alloy steel coltainin.g chromium. molybdenum, and 0.30% carbcn. 2-2.1.4 Stainless Steels All stainless stees contain at least 10.5% chrochro1s at least eellent nless Ai from which excellent corrosion resistance is obtained. Apparently, a very thin, transparrnt, and film of oxide forms upon the chromium surface. This film is inert, or passive, and does not react upon exposure to corrosive Materials. There are three broad types of stainless steels. as defined by thee
.
cf. content as the carbon difficulty. a result toof 0.30%, rpidmartensitic. some martcnsite may form as increases coling. ftheyare coyfole too arapidlt aftr ldi, Austenitic stainless steels, which have an austenicooling. If they Arc cooled too rapidly after welding. tic structure at rom temperature, arc known as the 3W series (AISI). These materials have excellent ducmedium- arid hil[.-carbon steels may harden, but pretilit, at heating to tM00 0F or post-heating to I 100'F will very low temperatures, the highest corrosion t remove britde nicrostructures. resistance of all steels, and the highest scale reThe yield strength of low-carbon steels is on the temperatures. strength at elevated itancc and steels is of high-cat bon psi, -while that be order of 46.0000,000 m achine, but steels are difficul, ustcnitic hncr A eelstioof fselastio canok modulua s ie toot4rt fom. psi.while psi.thtThe modulus 4of of 150,000 theorder order or. or the formd when carc is given to the r-te of workv miiiion for aii m in tension remains at 3S0 city ban steels. Core (Brinell) hardness ranges from 4 haideniiig. They art not harderiable by heatt rcatfrobon steels.Cove 40Brfnlhihardnessranges c rbom, ment. Welding is 'one best in an inert atmosphere, becae f the low thermal conductivity, care must be for low-carbon to 400 fo higher carbon. 2-2.1.3 Alloy Steels Alloy steels are those that contain significant amounts of such alloying metals as manganese, molybdenum, chromium, or nickel, which are added in order to obtain higher mechanical properties with heat treatmoent, especially in thick sections. A family of extra-high-strength, quenched, and tempered alboy steels has come into wide use because these c...a. rial.s havc y..ld s;reingths. of more than IArlflW psi.
The alloy steels have relatively good resistance to fracture, or tough-ess. Weldability is good, and machinability and castabihty are fair. The alloy steels generally can be hardeneC to a greater depth than can unalloyed steels with the same carbon content. Many of the alloy steels are available with added sulphur or lead for improved machinability. However, resulphurized and leaded steels are not recommended for highly stressed aircraft pa. ts because of drastic reductions in transverse properties. The alloy steels arce somewhat more difficult to forge than are the corresponding plain carbon steels, and the maximum recommended forging temperatures are about 50%
taken to avoid cracking. Carbide precipitation is minimized during welding by selecting one of the stabilized grades, e.g., 321 or 347. Ferritic stainless steels are magnetic and have good ductility. Because of the low carbon-to-chromiumn ,atio, the effects of thermal transformation are eliminated and the steels are not hardenable by heat treatment. They also do not work-harden to any grcuit e-tent. are machined easily,. and arc formed type readily. A general-purpose ferritic stainless is istp 430. Martensitic steels have a higher carbon-to-chroimiurn ratio and are hardenable by heat treatment. They are characteriLed by good ductility, hardness, and ability to hold an edge. These steels are. magnritic in all conditions, are tough and resistant to impact, and attain tensile strengths of up to 200,000 psi when hardened. Martensitic steels machine very well. Type 410 is the most widely used steel in this group. .,A
2-2.1.5 Precipitation Hardening Steels Precipitation hardening (PH) steels arc those that harden at relatively low temperatures due to the pre-
lower
Lipitalion of copper, aluminum, or titanium inter-
Cold-forming. it performed, is done in the annealed con6ition because of the high strength and limited duaility of heat-treated materials. Notch toughness of alloys in the heat-treated condition is much better than that of the carbon steels. Cor-
metallic compounds. They may be nonstainless or stainless. The best known is 17-4 P11, which is stainlcss by composition and is used fcr parts requiring high strength and good resistance to corrosion and oxidation at temperatures of up to 6 00F. 17-4 PH is
2-2
"
,•
:% "".
4,4
4-
.enerl-p
.
AM,,i /06 12 inartensitic in nature, but other precipitation hardening steels may be austenitic. Forming properties are .nuch the same as for stainless steels: forming must he accomnplished before heat treatment, and allowance must be made for the dimensional changes that occur during the hardening process. Strcrngth properties ate lowered by exposure to temperaturrs abosc 9750: for longer than 0.5 hr. The heat-treating procedures are specified in MIL-H-6875. •
--
2-2.1.6 Maraglng Steels The maraging steels are not treated in the refcrences given in par. 2-2.1.1; henc¢ they arc discussed in somewhat greater detail here. The term "maraging' is derived from the capability o! the material for age hardening in the martensitic condition. The distinguishing features of the 18% nickel maraging steels are that they arc designed to be martensitic upon cooling te room temperature after hot-working or annealing, and t- be agehardenailc to ultra high strengths in that condition. The 18% nickel maragirg steels essentially are wrought alloys. The nominal yield strengths of four "well-eitablished grades are 2Q., 2M), 3M).). and 35( ksi, The ability of these steels to transform into martensite upon cooling from elevated temperatures is im'paried by their nickel content. The transformation, which begins at about 310'F and ends at about 210°F, is of the diffusionless or shearing type. The formation of martensite in these steels is noi disturbed by varying the cooling rate within practicalec limits. Hence, section size is not a factor in the process of martensite formation, and the concepts of hardenability that dominate the technology of the quenched and tempered steels are not applicable with the maraging steels. The l8Ni maraging steels may be cut with a saw in the annealed or hot-worked condition. Alternatively, oxyacetylene and plasma arc torches may be used. Hot-rolled or annealed maraging steels can be sheared in much the same manner as can the quenched and tempered structural steels that have yield strengths in the vicinity of 110 ksi. In grinding, these steels behave in a manner similar to that of stainless steels, using a heavy-duty, water-soluble grinding fluid. The maraging steels can be hot-worked to finished "or semi-finished products by all of the standard methods of forming that are used for other steels. To avoid carburizing or sulfidizing, the metal should be free of oil, grease, and shop soil before heating. Fuel "withlow sulphur content is preferred. The meta! can be press- or hammer-forged at temperatures ranging 'from 23000 down to I500°F. Forging is completed at ý,"latively low temperatures. The objective is to refine
S~2-3
the grain struchtre, thereby enhancing ihe strenpth and toughness of the sleel. A minimum redution of 25'v. in thickness during the finil forging cycle is recommended to produce optimum mectianical propcrtices in the finished product. Ho, bending, hot drawing, and hot spinning are accomplished at 1500O- 180O F:. Cold-forrnirg operations are performed on the annealed material. Even in the annealed condition, the 18Ni maraging steels have yield strengths of up to 120 ksi, approximately four times those of deep-drawbody stock The tensile elongations of these steels in the form of annealed sheet may be as little as 3-4%. These factors impose limitations upon forming the sheet metal by tensile stresses. On the other hand, these steels work-harden very slowly, making them well suited to formirL methods dominated by shear. They can be cold-reduced by 80% or more, and shape. are formed readily by rolling or spinning. Flat-bottom cups can be deep-drawn to considerable depths. Roundtd shapes are formed mi re readily by means of the flexible die process. Cold-rolled, solution-annealed material is preferred. Rolling and welding of shwc, strip, and plate are cmm1on mithods of making cylindrical shapes. In the annealed condition, the l8Ni maraging steels are machined quite easily. In the age-hardened condition, machining is difficult because of the hardness imparted by the aging process. Although these steels have been welded by all of the common welding processes, the toughest welds are produced by the gas tungsten arc process or the electron beam kEB) process. For maximum toughness, the carbon, sulphur, silicon, phosphorus, and oxygen content must be kept at very low leveis. It is good practice to avoid prolongca times at elevated temperatures, not to preheat, to keep interpass ternperatures below about 250 *F, to use minimum weld energy input, and to avoid conditions causing slow cooling rates. Annealing is accomplished at 1500'F with air cooling. For improved combinations of stren~gth and toughness, the steel may be double-anne.aled. Thprocedure is to heat the material to !600°-l800'F, air cool to room temperature, reheat to 14000 -150O"F, and again air cool. Special furnarie atmosphe:-es art; required in order to prevent carburization, sullidation, or excessive oxidation. Age-hardening is accomplished at 9000)F, the time varying from 3 to 6 hr. Air is used commonly as the heat-treating atmosphere. It is advisable to maintain thc temperature at all part: of the load to within - 100 of the desired temperature. The nominal mechanical properties of the agehardened 18Ni tiaraging st~cts are listed in Table 2-I.
ca:~toas-4s thc: des~'icer in formulainig the re-
I~s~i
Adtona r~r~ daa h~h-t~cdgt
~
~
rnaraging btecis will be round a iv .I 2-.2.2 NON~FRUOUS MFITALS
o
d
E~nparativc
-
mcchfinical properties for mreprsenta-
lees3 ive 1aonluCW1_us alloys arc giken in Tabas- 2-2 Im;nIAly tlicii ols&.d Abrief review of nonferieus i-' af l Aus-a ý1 wco hk:,mprlrs kiusiv applicatkio to the cosru~m cur-wniirJ a as of some of [thc paramr.ters j,o,-vcrning *lýc Jioic; i.ofD~*{~ cus.ý. r of v'.-inum a~lcys, along with design at one metal amnoneg many for a a tý~la appficatiot,. smay1stdrdization ocz.uolacrIpe alonga with di-tail ticsigni Jita, is foune io AN11 Y nicnis, ir'cludiiig militury. fcdea ad, anid iindustry specilt00 arid in MIl.-1-DBI-5. ficatiowis. ThCSe Npe'.-ification, cover most of the uses Metais siuýh as alkmnrii¾m, in-gflesium, or iiof alulnihiaum in detail and shoutld IN consulted hefore rclativclv their tanium may be sailectcd beciuse of ein poadrgwt lilht weIghts. Other factors in material seteciion inailaminum allos are deexcept~or.s, fw Withi clectrica'i and thernial esstance. crrsin *eud nwogtpo o atn o o indete coraductivftty. lubricity, softness, cost and ease of, d-icts, but not for both. Althouth some genera!fabrcation. hardness. stiffness, and faoigpae resispurpose :-illoys are available. compositions rnoimally tance. UsAually, ii is the sum of a number of factors formulated so as to satisfy tspecific rqirements. arc of that in',1uences a designer to select the sequence mnoic %&idely used And readily avaiwable copiThe a constitute that processes materials and fabrication positions. are covered by G(.?vcinment specifications. design item. This discussion is inteiided to provide Most are adaptable to a variety of applications. The Aluminum Association has devised afor TABLE2-1 digit system -for wrought alloys in which the first 110*v fABL12N, ig% no " ý. A number designazes tnhe major anulloyi ungemiciii. Ti u.N. MARAGING ST EELS I is pure alu-minum, 'A is copper, 3 is manganese, 4 is
2-2.2.
*
Ertc
I
silicon, 5 is magnesium, 6 is magnesium and silicon, and V is zinc. The last two digits are supposed to ULTIMATE TE14SILf 00 designate the aluminum purity, but the exceptions 365. 00 294,000 260,000 210,000 NGTH,psi ST RE destroy. the rule. hlowever, the more frequently used 4 ~ 55,001290,000 255,000 000 SRENTH, si 0.2-YIED become familiar to the designer. The 0.~YILD * TREGTHPSI355000alloys aluminum casting alloys usually are identified by ar125 13.1 11.8 100 ELONATIN,-. ELN A IN,,IU0 T 10 1 . bitrarilv selected com mercial dicsignations of twoand three-digit numbers 2.0 570 6.0 0.0I OF~RE~o * IOREIJC NOT UCHTENSIL STREA' N50 0 6!. 2. Most alumninum alloys used for wrought products contain less than 7% of alloying elements. By regula33 qu00 420nn 1 190m00 I25.000 ~ STEGH ~ TE0IL NOTCH 19.0'0' -t tion of the amounts and types of elements added, the operties of the alumninum can be enhanced and its pr.0 17.019. 1. CHARFY V-NOTCHII-Is working characteristics improved. Special comLITIGU (0DUCYCLES 1000 2,00 115,000 115"U000 positions have been developed for particular fabrication processes, such as forging and extrusion. Wrought alloys are produced in both heat-treatable 1 75 60 55 HRDNCKESLS'~ HARDNESS0 nn nonheat-Lreatable types. The mechanical properties of the nonhecat-treatable materials may be varied 220026,0 8,0 siE1 STRENI~H, YIELDOPE
FSLRIES
-250
-30130
_
-2,
-------
-~
--
-
-
!
I
-
--. ------------
'
-
strain-hardening or by a combination of strain-
TABLE 24by
COMPARr
TABLECIA 2-2ERIE A
h.irdering and annealing&...
I he aluminum alloys specified for casting purposes contain one or more alloying! elements; the maximum . .n~u of any one element must not exceed 12%. designed Sonealoy are deindfor usc in the as-cast con-
COMPR~n E MCHAICALPROERTES FOR SELECTED NONFERROUS, ALLOYS PTY FROPE _
--
_-A
T
A29.C 14
1% STRENGTH. LL
IENSILL STRENGTH,.~
-.-ý
34
ELONGATION,
MOOIQ"US. 10Il
'Wo
Llui
6.5
2017
VANSM 'cr[oun:.
I C-i
I4A,~~IVCR
ition; others are designed to he heat-treated in order to improve their mechanical properties ;nd dimensional stability. High strength with good ductility can be o~btained by selecting the appropriate comnpo-
32 -
17
___
12
5
66
10.4
5.5
-
230 -HANUNESSIB ..
2-4
ndha
-
retet
~sitionanhetramn.
AMCP 706-202
* . S"
-
_. TN•
S,)
The hcat-trcatment and temper desinations for aluminum arc long and complex. The desinations most frequently stamped on products are: F-as fabricated; 0-annealed; H-strain-hardened (many subdivisions); T2- (cast products only). T4-solution lihat-treated and naturally aged. and T6-solution heat-treated and artificially aged. The heat trcstmcnt of aluminum alloys is detailed in MIL-H-6083. The processes commonly used are solution heat treatment. precipitation hardening, and annealing. A .small amount of cold-working aftcr solution heat "treatment produces a substantial increase in yield strength, some increase in tensile strength, and some loss in ductility. Rapid quenching will provide maximum corrosion resistance, while a slower quench-used for heavy sections and large forgings-tcnds to minimize cracking and distortion. Momt forming of aluminum is done cold. The ternperature chosen permits the completion of the fabrication without the necessity for any intermediate annealing. Hot-forming of aiuminum usually is perS formed at temperatures of 300°-400°F, and heating periods are limited to 15-30 min. When nonhcatitreatable alloys are to be formed, the tempef bhuuld be just soft enough so as to permit the required bend radius or draw depth. When heat-ireatable alloys are used, the shape should govern the alloy selected arnd its temper. To a great extent, the choice of an alloy for casting is governed by the type of mold to be employed. In turn, the type of mold is determined by factors such as intricacy of design, size, cross section, tolerance, surface finish, and the number of castings to be produced. In all casting processes, alloys with high silicon ,-onte~nt ee m~fmi in thwprndctinmi of narts with thin walls and intricate design.
mizt distortion due to expansion and contraction. Molten aluminum bsorbs hydrogen easily, and this may cause porosity during cooling. Because they provide a protective inert-gas shield. TIG and MIG welding are common choices. TIG is an incrt-gas shield-arc process with a tungsten electrode, and MIG is an inert-gas, shielded-metal-arc process using covered electrodes. A suitable flux, and mechanical (stainlcss steel brush) removal of the oxide film just prior to welding. are mandatory. Certain aluminum alloys - 2014, 7075, etc. - arc extremely difficult to fusion weld (excluding spot welding) and normally would not be used in structural applications when welded. Brazing is somewhat more difficult, and soldering of aluminum ir extremely difficult. The other joining processes include riveting and adhesive bending, both of which are used extensively in aircraft structures. Applications for aluminum in helicopters include the 4heet-metal exterior surface of the fuselage. framing, stringers, beams, tubing, and other usages where the density, corrosion resistance, and ease of fabrication of alum,'ium give it an advantage ovcr steel and where its high,, star-eg•,,sth, and oadulus properties give it an advantage over magnesium. 2-?.2.3 MNgneslum Alloys MlL-HDBK-693 provides a comprehensive discussion of magnesium alloys and their properties, and also describes design, fabrication, and performance data. Numerous Military and Federal Spccifications covering specific shapes, forms, and p ,,:esses also are summarized. I ,e outstanding characteristic of magnesium is its igl.i weight. This is important in helicopter design,. ,,-•. payvload ratio is a direct function of vehicle ,.,h Magnesium is two-thirds as hevy as atumi..
The most easily machined aluminum alloy is 2011T3, referred to as the frce-cutting alloy. In general. alloys containing copper, zinc, and magnesium as the principal added constituents are machined the most readily. Wrought alloys that have been heat-trcated have fair to good machining qualities. The welding of many aluminum alloys is common, practice because it is fast, easy, and relatively inexpensive. Welding is usclul especially for making ',akproof joints in thick or thin metal, and the r,. - ess c€n be employed w;th either cast or wrought. :,num or with a combination of both. The re' -:;.i. low melting point, the high thermal conductiviy , the high thermal expansion pose problems. •., heating is necessary when welding heavy secti,:. otherwise, the mass of the parent metal will cond,. the heat away too rapidly for effective welding. , rapid welding process is preferred in order to mini-
rm,;w
.1one-fourth as heavy as steel. The low densi• tI effective in relatively thick castings, where ,s :;r -sed rigidity of magnesium is an additional hcnefit or this reason magnesium is used freq••iv., . main rotor gearboxes, motor trans,,:..sings, and many other load-bearing :ilphpii .' .mn helicopters. Most of the helicopter pm C .: . have several hundred pounds of maincsiw, . construction. The ,--.: (American Society for Testing Mcr •1 .-: '•it nclature system is used exclusively in u',: , . ,..,sium alloys. In this sytem, the first ,t I. . .1' the principal alloy elements, while th, n -. ,ate the rcspectivi, percentages. I'. ., .... aluminum, F i -. earth, H •,, . .:.,, ;ým. 1. lithium, M manganese, Q .iye I . -,o ,.'inc. By this designation, AZ91C - Co..i .. i.ioy of magnesium containing 9% i\
-.
..
"'*
'A''
j'. "'
`
AMCP 706-2012 r
.luriiinum. V; zinc, and having a "C" variation.
The heat-treat and temper designations for magnesium virtually are identical to those for aluminum. -:1 he temp~er designations used urc those in ASTM 8296. There are four groups of mugnesium casting alloys. The Mg-A and Mg-Z binary systems are tiesigned Iir use at temperatures belo%% 300'1[ and are ~~iof Ios~er cost. The Mg-F and the Mg-Fl binary s~stents are designed for good strength in the 500'0 8004:~ range. The choice of casting composition is dictated largely by certain features of the design, and by cost and irecthod of production. For magnesium alloys, the important casting proce~ses are sand, per. mnciaret miold, and die. The choice of a casting proc-ess depends upon the size, shape, and minimum seclion thickness of the part, and upon the tolerances, types of surface finish, number of pieces to be produced. and relative cost of finishing the part. Magnesium alloys, both cast and wrought. haie outstanding macmiinability. Greater depths of cut and higher cutting rates can be used with these metals than with other structurai metals. Magnesium does slt
?a
%%,U .
s
U
94
ns
;ý
siderahh higther tenmperatures than aluminum often gise it advantages bor partic:ular applications, as in the hot structures and exhaust ducting for helicopter power s)-steins. Indeed. increased paj loads resulting front %% eight saving~i catmr ~a fseth initial costs, and in the long run titanium may prove less costls for seii plctosta okrpie mtaterials.spih aplcto hnlserrid 1Itnmisaihlt mlc ilsivrou srought shapes and in at side range of' alloyed ane unalloyed grades including billet. bar, extrusions, plate. iheet, and tuh.*ng. The mill products can be grouped into three categories according to the predomninant phase in their inicrostruclure: Alpha, Alpha-Beta. and Beta titanium. There is no single acceptt'd system for the designation or classification of titanium and its alloys as there ure for other metals. Titanium iactual", is easier to machine than the stainless steels becauhe the effects of work -harden ing are far less pronounced. Titaniumi requires low shearing forces, and is not noi-.h-sensitise. Because of these properties, it can he machined to extremely low micro-inch finishes. On the other band, the sharp
,*M4
-
-
-
the shearing point to heat rapidly. At elevated temperatures tItnmtedtoislvayhig within contact, and the cutting tool is dulled revdily. F~urther. the cirhides and oxides oin the forged pie.es are extremecly abrasive to tou's and miust be removed by tiltric-hydrolfluoric acid treatment prior to machining. Osciall, considerable kno%%hoA is required for the economical machining of titanium. Titanium assemblies are joined by spot, scant, flash, and pressure welding technique-.. In fusion welding, the TIG process is used-. heavy welding also requires Incrt g.. k.
other metals, and welding of magnesiam to magnesium can be accomplishmed reliably only by a skilled operator. The metal also cannot be soldered. properly. Thus, electron beam (EB) welding is the mostsatsfator wedin prces, athogh luxdip mostsatsfator wedin prces, athogh iuxdip brazing also may be used; care must be employed in removing all of the flux because of the danger of corrosion. The best method ofjoir .ig magnesium in thin sectonsis ahesve ondng.mal y
welding is quite saiLffactory. 2-2.2.5 Copper and Copper Alloys
2-2.2.4 'Titansium Alloys MIL-HDIIK-697 contains a comprehensive description of titanium alloys arnd their properties, and discusses design, fabrication, and performance. In addition, seven Military Specifications for specific forms of titanium will be found in Refs. 2 and 3. Although titanium is relatively costly, its high strength-to-weight ratio, excellent corrosion. resistance, and capability of performing at con-
tageous in inserts, studs, bushings. etc., where owk load is desired. Beryllium copper is useful for sprink's and oiler applications where its good modulus, hardness, fatigue resistance, and ease of formiing are advantageous. tlo%%e~er, copper alloys are the hcaviest of the common structural metals, and, therefore, have at %%eight disadvantage in aiiborne applications. The various types of copper and its alloys are
2-6
.
.
I! 1* A
Ionn. C;11;%-*
polish the material in order to obtain an. extremnely fine finish. The chips from machining readily clear ihe work and the tools. Because of its position in the electromiotive series, magnesium is subject to more corrosion than are the other structural metals. The many corrosion problerns associated with the use of magnesium severely limit its use in rotary aircraft. Magnesium alloys shall not be used for parts that are not readily accessible for inspection, application of protective finish, and replacement. magnc.ýiuin cannot be wcided satisiucturily to
A 1;oniprehensivc discussion of coppeF and copper alloys and their prtop-rties. design and fabrication characteristics, and design and performance data is contained in MIL-l-ll)BK-698. h aiu om fcpe n otraly have found only limlited use in helicopters Their therand electrical conductivity properties are ad.-an-urct
.
tl -A
1
*I
AMCP 705-202
hetter known hy namec than by code number 1 lie' lerin copper is used when the material cxceds 99.4'4 purnty. The principal alloying agent of bra... is zinc, v.hile tin isthe principal alloying agnin in bronie. Thc beryllium coppers have sinall percentajges of herylhum,. producing
u remarkably hard. hiph-modulus.
high-strength, inonsparking mat erial. The copper and copp~r alloys arc ca.st readily in all of the various castling processes. The allo.ýs are coldformed catsly, and are capahleof being rolled, drawn. spun, and flanged. In hot-%orking they arc rolled, cxtruded. pierced, and forged. The machinability ofcopper alloys is excellent. For sand castings, low speeds and coarse feed-- are used for remosing the scale in order ito increast tool life. It is better ito remove the scale by sand blasting and pickling, The copper alloys are %elded readily by all the sselding processes, although their high thermal conductivit) i%a problemi. The,, arc adaptable ito brating and are the eiasiest of ;ill metals to solder. 2-2. FL:cTOI.Y~c (1IN OFfl~;SIII2-2.3 MEITALST AmeO O 1SM1Dissimilar metals, as defined in MIL-STD.889. should not be used together in helicopter applicathc m ating surfaces are insu lated adetinis unless When tape is used between two dissimilar . qt\ quately. S metals, such as in the mounting of a magnesium gearbox toa~n adurnintm airframe, the contractor miust insure that there will be no loss oi- mounting torque as a result of normal usage and vibrations. Metali can be grouped in four categories, as shown in Table 2-3 Metals grouped in any one of the calegorits in Table 2-3 can rye considered similar to one another, while those metals placed in different groups should be considered dissimilar to one -unoher. The categorization does not apply to fasteners - such as rivets, bolts, nuts, and washers -- that are comrponent parts of assemblies and usually are painted prior ito being used. Instead, the metels referred to arc surface mnetals. F-or example, zinc covers all zinc parts, ineluding castings and zinc-coated paits. TABLE 2-3 GROUPING OF METALS AND ALLOYS (M IL-STD-889)
~~McGraw-tfill,
GFUU' 5I'M AD I Al UM'I.1UKIAt OYS !.",2 Ei`,t, 5351, (;ROUI' 11
IlI 0i~lDUI' GRUPI
6061 AND 1410
IllR At 101 t'ADMIiIM PNI . ANO All A'.~lI AND (INCIIDIND THElAIDIMItRIIK All CIYSIrl GllILIr' 1,
IRN I EAC!).ANU TINANDLHIR~l All (,i lEXlCu'il AIlNI 1SS SIFELS) C01111. I'll MILIM NICKEL Sit VIER(,OLD. P1MINUM TITANIUM. COBA, AND 1`4ODIlUM AND 1110111 AllIOYS SIAINI.LSS STEEl S AtOL itlAP111fI
______
Par. 2-2.2.2 details the use of alumninum alloys in 11clicopter construction. Aluminum alloys used in helicoýptcrs may contain copper ot zinc as an cssent ial const ituent. In sheet form, thesecalloys rcsusceptihlc ito corrosive action resulting in a loss of strength of the material, which becomes brittle withoiut evidence of surface change. Aluminum alloys containing magnecsium, magnesium and silicon. and chomniumn as the essential alloying constituents arc mnuch more stable under prolonged weathering condition% than are the aluminum alloy% containing copper or zun::.4 Cadmium behaves similarly to zinc as a coating mectal hi.-affording electrochemical piotection of ferroins%metals against corrosion. Cadmium plating thus can he: used to put ferrous metals into the same group a%5 aluminum al~oys. giving them a similarity. A dctailed ditscutsion of coating processes can he found in p~ar- 2-tv. In general, when two dissimilar metal surfaces come in contact with one anothcr, a corrosive action called galvanic action can take place. Coating iitals are used as thin layers between dissimilar mietals to prevent this type of corrosion. Table 2-4 illustrates the position of metals with ~.I r fth V lack of su ce ti.lt i.h__r red t to galvanic action.
2-3
1
NONMETALLIC MATERIALS
2-3.1 GENERAL This paragraph discusses the applications of the thermoplasstic and thermosetting plastics, elastomners, fabrics, and transparent maiteials. Other materials - such as glass in light bulbs or optical piping, ceramics and mica in electrical insulation, and cairbon and graphite in lubrication or electrical contacts - also play significant roles in helicopter construction. The nonmetallic materials used in composite structures, reinforced plastics. and other composite mnaterial% are treated in par. 2-4; plastic materials used as sealants and adhesive-; arc covered in par. 2-5. Comprehensive discussions and detu~led design data will be found iii existing documents. Among these are: MIL-H-DBK-700, MIL-HDBK-l7, the Modern Plastics Encyclopedia, published annually by and the Malerials -Selector Isjxuv, published annually by Reinhold Publishing Co. The major disadvantage of plastic's is their low moduitis. which is in the order of a few hundred thousand psi compared to 10 or mome million psi for metals. Thiey also arc more- sensitive to heat, soften-
ing markedly at 400~0 F and below. On the other hand.
plastics can he as strong as steel, be lighter than magnesiurn, aind have better abrasion resistance than
~2-7
.
I
4X metals. Normal corrosion is not a problem. Although they are nonconductors for electricity and poor conductors for heat. they can be exceedingly tough and wear-resistant, and can be fabricated in a variety of waya. When judiciously selected and pioperly used, they often can p.rform better at lower cost than any other material. TABLE24 POSITION OF METALS IN THE GALVAN'C SERIFS
MAGNESIUM MAGNESIUM ALLOY
shapes. plate. sheet, and film and in a wide range of shapes. sizes, and thicknesses. The materials are machined, or shaped readily by thermoforming proccsscs. Many items may be purchased in thc finished form as produced by extrusion or injrction molding: included arc screws, nuts. bolts, inserts, grommets, straps. pins, knobs, handles. instrument facings. housings, boxes, conduits. electrical receptacles, covers, rails, runners, guides, snaps, and slides. Many of these items are supplied as off-the-shelf inventories in a variety of sizes. Nylons and polycarbonates are known as the enginecring plastics. Nylon, a polyimide. has high strength and high elongation, giving it a toughness that many applications depend upon. It has high
ZINC ALUMINUM 1100
modulus in flexure, good impact strength, a low coefficient of friction, and high abrasion resistadcC, as
CADMIUM ALUMINUM 2017 STEEL OR IRON CAST IRON LEAD-TIN SOLDERS
well as good fatigue resistance under vibration conditions. Its primary disadvantages, though not significant, are dimensional change with moisture absorption, and the need for incorporating carbon black in order to )rotect against ultraviolet degradation in outdoor use. Nylon is used in gears, arms and other contact applications, and in pressure
CORRODED END ANODIC.1
SLE)
•
LEAD
:I
TIN -•-;T
.
;
tubing, belting, and we.ar pads. The polycarbonates are aromatic esters of car-
BRONZE
COPPER-NICKEL ALLOYS TITANIUM MONEL SILVER SOLDER NICKEL INCONEL CHROMIUM-IRON 18-8 STAINLESS 18-8-3 STAINLESS *.SILVER
,• L
GRAPHITE GOLD PLATINUM PLATINUM
______
PROTECTED END (CATHODIC]
OR MOST NOBLE)
SOURCE: REFERENCE DATA FOR RADIO ENGINEERS
FEDERAL TELEFHONE & RADIO Co. 3RD
2-3.2 THERMOPLASTIC MATERIALS Thermoplastic materials are those that soften when heated and harden when cooled. Typical of the thermoplazic family are the polyvinyls, acrylics, nylons, polycarbon tes, aitd fluorocarbons- Often, these have linear micromolecular structures. Products of these materials usually are formed by extrusion or by inJection molding, and they are available for manufacturing in the form of rods, tubes, contoured 2-8
boric acid. They have excellent rigidity and toughness. high impact strength, and low water absorption. They are stable dimensionally under a wide range of conditions, are cieep-resistant, and are transparent and stable in sunlight. Probably their major deficiency is that their fatigue resistance is lower than is desirable. Polycarbonates are used in shields, lenses, ammunition chutes, knobs, handles, etc. The acrylic of interest here is polymethylmethacrylate, better known as Plexiglas. This plastic has crystal clarity, outstanding weatherability in optical p, operties and appearance, dimensional stability, good impact resistance, and a low water absorption
.4.
rate. Its major deficiency is its low resistance to
'-
scratching. Its major use is as window glazing and for such applications
as transparent aircraft covers;
covers for signal lights, where its ease of coloring is advantageous; and in other optical and instrumentation applications. Its use as a window material is discussed in par. 2-3.5. For helicopters, the polyvinyls are used largely in the form of sheeting simulating leather or upholstery fabric. These are very tough and wear-resistant. In the transparent form, they are used to make pockets and holders for documents and maps. The fluorocarbon poymers have excellent thermal stability at continuous temperatures of 400°-550*F.
*
.-
-
AMCP 706-202
)
They virtually are inert to chemical attack, have exccllent damping properties, and have outstanding electrical characteristics. such as high dielectric strength, low dissipation factor and radio frequency (RF) transparency They arc used widely in microwave components and high-frequency connectors, as well as in wire coatings, gaskets, and electrical tcr•minals. 2-3.3 THERMOSETTING MATERIALS H M TT MT ASfusible. A lthou!h there is a great diversity in the chemical mak:up of thermosetting resins, they have one chaIacteristic in common: once they are cross-linked, tl-ey do not soften undoi heat and cannot be formed by thermofotming processes. With the application of heat, thermosetting resins undergo a series of changes that are irreversible. The polymeritation reaction that occurs results in such a high degree of crosslinking that the cured product essentially is one molecule. In many cases, this results in a highly rigid molecule of good thermal stability. The thermosetting rcsirts usually are used with fillers and rcinforcement. Three of the most widel. used of these materials are the epoxy. phenolic. anid polyc~tei iesius. Thcsc are employed extensively with Fiberglas fabric. %%ith chopped fiber in laminates, in sprayed forms, in filamnent-wound structures, in honeycomb sandwich structures, and in combinations with halsa wood or formed shapes. The epoxy resins arc based upon the reactivity of the epoxide ktroup and generally are produced from bisphenol-A and cpichloiohydrin. Epoxies have a broad capability for blending properties through resin systems, fillers, and additives. Formulations can ne so.i armnex1t.'01c or F-11,U T'are.. ... able as prepolyrners for final polymeritation in the form of powNders and liquids with a %ide range of viscosities, Some cure at room temperature, while others require curing at elevated temperatures. The powders may be transfcr-qilded byi machine, and the liqiids may he cast. Mure-oftcn, the liquid is used to inipregnate materials for bonding. The outstanding characteristic of epoxies is their capability to form a strong bond with almost an) surface. Fur this reason. they are used widely in adhesive formulations. The molded produtcts have high dimensional stability over a %ide range of temperatures and hunridities, excellent inmchanical and shock resistance, good retention of properties at 50091-, and excellen! electrical properties, The phenolics are the oldest and the least cxpensive of the plastics. 1-thermosetting he basic resin is ianufactured by weans of a rc,actiorr bet ncii
phenol and formaldehyde. Thi. resin is blended with dye, filler, and curing agents to make the molding powder, which is called the "A" stage powder. Pokkders such as these are molded for 2 min at '25'F at 15(X) psi pressure. As the granules are warmed by the hot mold, the resin melts; the material flows and fills the cavity, further reacting and going through a rubbery "B' stage. With further cross-linking it reaches the -'C" stage, at which it is hard and inFillers used in typical phenolic molding powders are fibrous in nature; their interlocking fibers act to reduce the brittleness of the cured resin. Wood flour is used most commonly, while asbestos and graphite fibers form the i.onventional heat-resistant plastics. Paper and fabric fillers are used for high impact or shock-resistant phenolics. When the powders are used for lamination or in making composite structurc-:. a solution of the resin in alcohol is used to impregnate the fabric and then "B" st.ged. The layers ot" impregnated fabric are laid together and then cured b) heat and pressure. The advantages of phenolics are their low material processing costs, dimensional stability. excellent hrc ktia fpris 1 lodb1rn teristics, and good weathering properties. They are used in electrical components, receptacles, conduits. housings, etc. IThe polyesters are plastics formed of chains produccd by repeating units of a polyacid and a polyglycol. They may be aliphatic or aromatic. Familiar forms - fibers such as Dacron synthetic fiber or Mylar film - are the so-called linear polye.;ters. More important to applications in helicopters are the thermosetting resins. These arc the three-dimensional or cro.-linked polyesters that are formed by bridging ,i unsaturated polyesters. In this form, the polyester is supplied as a syrupy liquid that - %hen mixed with a small amount of curing agent, ,pplied to a fabric, chopped Fiberglas, or filament tow, and laid over a form - rapidly reacts so as to establish a rigid structure. Such structures are of particular use for ra. domes because of their RF transmission and excellent weaaherability. They have high modulus and impact strength as well as excellent flexural and tensile properties. The polyesters may be cast in order to produce glazing materials. Another highly cross-linked family of polymers consists of the urethanes. 1 hese arc formed by the reaction ot isocyanates with esters and unsaturatcs. In the process, carbon dioxide is evolved and forms a highly porous structure. "[he stiffness ranges from soft, flexible foams to highly rigid foams. The flcxible foams are employed for cushioning and padding 2-9
.
-
1xvln
SECP 706202 .and to reduce noise and shock, as well as for thermal insulation. The rigid foams are used as light-weight stiffeners in structures.
MIL-HDBK-149 presents a comprehensive discussion of the technology of the elastomeric materials and their applications. From the standpoint of durability and performance, natural rubber remains in demand, and substantial quantities arc used in blends with SBR, butyl, and other synthetic rubbers. Natural rubber is a stcrospecific polymer of isoprene. Its applications in pneumatic tires, bumpers, shock absorbers, etc., as well as in belting, gaskets, and seals, arc well known, Substantial quantities con"tinue tc. be employed in hdicopters. Carbon black constitutes about 50% of the w,:ight of il-csc corpositions. A more advanced synthetic rubber is nt-nnr%,,e, a general-purpose synthetic made by emulsion polymerization of chloroprene. A notable characteristic ":t rabte•ri e ris i,.,o gasoline. foles, monrieating oils, and othur solvents, a d its excellent resistance to weather-oxidations, ozone, and ultraviolet light. It has goon tensile strength, tear resislance, abrasion resistance, and rebound character;stics, and excellent adherence to metal and fabrics. It provides average insulation and has excellent dicirctric strength. In helicopters, it is used to coat radomes and the leading edges of the rotors for protection against abrasion by rain and dust. It also is used in boots on other leading edges and areas where wear is a factor; and in transmission belts, hoses, liEcs sas,
&-1 ecica
apl-U----------
Another important family of clastonteric materials is the silicones, which are used in many diverse and seemingly unrelated applications. The silicones are organo-poly-siloxanes, having alternating silicon and oxygen atoms in the backbone of the chain. The silicone resins may be cast, extruded, or injectionmolded so as to form shaped products. They are available in sheet or bulk form; as a range of pastes and liquids for use as adhesives, sealants, and coatings; and as powders for foaming. They are stable continuously at temperatures from - 140' to +600 0 F, and initermittently to 70001. They are weather-resistant, hove high dielectric strength and a low dissipation factor, and are bonded easily to metals, ceramics, and plastics substrates. Aromatic solvents and chlorocoinpounds swell silicones, and they have higher gas permeability than do other rubbert. They are used as foaming agents, encapsulatin mesins, sealants, and in electrical applications, 2-10
2-3.5 WINDOW MATERIALS Gazing materials and methods of attachment are , discussed in detail in MIL-HDBK-17. That document also lists additional Military Spccification" coveting specific glazing materials, resins, cement., and proc-.sscs for the design and fabrication of widow systems. The optical properties of greatest significancefor aircraft glazing are surface reflection, index 9 'f refraction, absorption of light, and transmission of an undistorted image. The thermal properties of primary concern are the coefficient of expansion, the thermal conductivity, and the distortion temperatare. The major physical properties are density and hardness, or scratch-resistance; the major mechanical properties are tensile and compressive strength and the modulus of elasticity. The ideal glazing will be strong enough to withstand structural and operational (wind and water) loads, hard enough to remain unscratched, optically clear after a life of operatiun, unchanged by thermal loads, and unaffccted by the weather. Although no material- passess all of these desirable characteristics, there are several that perform ver) well. The three glazing materials that are employed most often arc glass, cast polyester, and cast aciylic (methylmethacrylate). Polycarbonate, an otherwise strong contender, has not yet been produced economically in large sheets with the required optical properties. Monolithic glas- is used in helicopters only when use temperatures ex,.ecd the performance t(mperatures of the laminated glasses and the poiymeric materials. The lamination of glass with plastit imt provcE. the -cs~s!_i. to- therma and rne-har.ca! stresses, and minimize; the possibility of complete failure of a panel, Splintering of the g'asi is prevented, although the load-carrying capacity of the laminated glass is less than that of plate glass. The plastic interlayer is sel,.cted so as to provide the greatest ability to absorb impact energy. Polyvinyl butyral is the most common interlayer material for both glass and plastic laminates. A new thermosetting, polyester-base, transparent sheet material has becn developed under the trade niame "Sierracin 880". It can be used for aircraft endosuren. that operate at suifitce temperatures of up to 300°F. and is characterized by its two-stage cure. After formiig and post-ctri~ag, the ultimate physical properties of this materia! are obtained. Sierracin 880 generally is used as a laminate with acrylic, and is described in MIL-P-8257. The glazing materiaS u;sed nast widely for helicopters is cast polymethylmethacrylate. In many air-
I
S" craft, it constitutes a major portion of the fuselage walls. For window applications, the stretched, modifled acrylic sheet is preferred, per MIL-P-25690. The nmodified material has slightly higher heat resistance than does heat-resistant polymethylmethacrylatc, along with better resistance to crazing and solvents. When stretched to W-I100% biaxially or multiaxially. acrylic sheets develop increased resistance to crazing, higher impact strength, and improved rcsistance to crack propagation - without detrimental effects upon their other properties except for reduced abrasion rFsistance and laminar tensile and shear strengths. The sheets may be formed thermally to diffcrent contours. Laminated plastic glazing materials are made by bonding two or more layers of acrylic or polyester plagtic sheet to a soft plastic interlayer by means of an adhesive. This process greatly improves the impact and structural strengths of the material, Laminated plastic glazing materials arc defined in MIL-1N5374. Differing thermal expansion rates of glazing materials, edge attachment mmieriais, and metal air-
frames present one of the major probicms
iM I
design of window-glazing. For all types of glazing, an edge wrap is used in order to minimize the propagation of stresses originating in cracks and chips at the edges. The preferred wrap is two or more layers of polyester (Dacron) fabric, ,,oven from twisted yarns and impitgnated and bonded in ?lace with epoxy resin. Th wrap overlaps sufficiently on the glaze material and extends sufficiently beyond the edges to absorb the stresses of aitachment closure and at the same time "o distribute the load uniformly across the window. There are many closure designs, but the prefcrred enclosure will be designed so as to hold the glaze securely in a sliding grip ir. such a manner as to allow for reciprocal longitudinal motion - as the glaze expands and contracts - while always applying a comprehensive load endwise. This may be achieved by placing a compressible, neoprene-imprcgnated tube at the bottom of the closure channel. The closure will be attached rigidly to the airframe. The contacting areas between the closure and the edge wrap will be sealed with a flexible sealant, prefe-rably one made of silicone,
-
24 COMPOSITE STRUCTURES 2-4,1 FIBERGLAS LAMINA'TES *
".
Of all the fibers available for the rcinforcement of plastics, glass is used by far the most widely. Of the various glass compositions, only two are important in "aircraftconstruction: "E" and "S" glass. "'E' glass is used extensivily; "S" glass provides greater tensile
A CF 706.a01
strength and modulus, but is considerably more cxpensive. The major advantages of glass-reinforced plastic (GRP) over isotropic structural mate•ials (primary metalsý include: I. Formability and versatility. Large complex parts and very short production runs are practical. Because there are few limitations on size, shape, and number of parts, design freedom is maximized. In addition, the reinforcement can be oriented as desired in order to increase properties ir. specific directions. 2. Chemical stability. GRP is resistant to most chemicals, and does not rust or corrode. 3. Toughness. Good impact resistance is a feature of GRF. 4. Strength-to-weight ratio. Specific strength of GRP is very high. For example. unidirectional GRP has a specific strength about five times that of the commonly used steel and aluminum alloys. 5. Insulation. GRP is a good thermal and electrical insulator, and therefore, wiil transmit radar and radio waves. Wk . ......... , ..... r .. .atched n he readily and effectively. On the other hand, GRP has certain disadantages compared to other construction materials, namely: I. Nonuniformity. Variations in material properties within a part and from part to part are inherent in most of the fabrication techniques. 2. Low modulus. Stiffness of GRP is relatively low. 3. Slow fabrication. Production rates are low in comparison with most metal-formirg operations. Thus, GRP coiitiUctmon is mnostudan'tgeou•s ar, parts with complex shapes that would be difficult to form from metal, foe parts with anisotropic strength requirements, or in applications where the conductivity or poor dent or corrosion resistance of metals present a probl'm. Sonic typical GRP applications in helicopter construction arc in canopies, covers, 'tad shrouds (for formability, specific strength, dent resistance); rotor blades (for formability, anisotropic strength, and stiffness)- control surfaces (for anisotropic properties, dent resistance, repairability); and antenna
housings (for radio frequency transparency, formability). It is conceivable that an entire helicopter airframc can be constructed from GRP, as has been
done o, several smal, fixed-wing aircraft. 2-4.1.1 Design Colsiderations Design rules and procedui ,. for reinforced plastics do not differ markedly from those for metals. Stress st:ain curves, however, are similar to those for wood 2-11
.
. ,
.
(
AreP 7062D2 (which also is a riber-reinforced composite) in that there is no yicld point. As Part I of MiL.IIDBK-17 contains conlsidcrable prapcrty data on specific materials, only generalities arc considered herce, ~ar hertw cscntal igreiens i glss. 7~reinforced plastics: glass fibers and resin. A finish, or coupling agent. that enhances adhesion between ile glass and resin usu'lly is used as a cobting on the glass, and may be considered as a third component or as part of the reinforcement. The resin system ha roe ~ detrmiingthe geneall th limtin chemicalyhathermlmiand e oletica pdopertiesiofgltm catemwicl, thertyealon, elcrclpoprtientto and of themi reinforcement predominate in determiining the basic --
mechanical poets.esters, propeties.laytips 2-4.1.2 Resin Systems
rtint
*
.
Essentially all GRP laminates arc madc with thcrmosetting re-sins that, when mixed with suitable catalysts or curing agents, arc pormanently converted to the solid state. Reinforced thermoplastics (RTP). whicn contain snort glaass ircrb. arc a rapiduy growing element of the injctlion-molaing industry: such parts. however, are not consider.-d laminates, Probably 95% of all GRP laminates are made from polyester, epoxy, or phenolic resins. For very-high. temperature service (above 50OOT), silicone (MIL-R25506 and MIL-P.25S18) and polyimide resins are available. These, however, have no known applications in current helicopter technology. 2-4.1.2.1 Polyesters These are by far the most widely used resins when thob eniirp rGP intisictry ic -nn-ijere-d. They are !ow, in cost, easily processed, and extremely versatile. Available types range from rigid to flChible: there are also grades thait are fire-retardant. ulhrav'ioletresistant, and highly chemical-resistant. The upper temperature limit for long-term operation of generalpurpose grades is 200*F, although temperature-resisresins are available that are useful up ito 5000F. These can be formulated for rapid curing at room temperature or with long pot-life for curing at elevated temperature. rhus, they commonly arc used for wet layups but prcpregs also are used frequently. Prepregs that cure by ultraviolet light also are available. Disadv:tntages of polyesters include high shrinkage * during cui-c, inhcrently tacky surfact. ii cured in the
presence of air, odor, and fire hazard in wet layup fabrication from styrene monomecr and peroxide catalysts. Requirements for general-purpose polyester larninating resins and lamuinates are contained in M IL-R7575 and L-P-383, respectively; fire-resistant resins 2-12
und lainin.,tes are detailed in MIL-R-215042 and Ml L-IP-25395. 2-4.1.2.2 Epoxies Epoxies probably arc thec resins most frequently used ior atircraft GRP laininates. Although about twice as costly as polyesters or phenolics, epoxy plctos rsn tl jeiepniefrms stapiatint. hcia-rsivefrm rechansticl, 4elietrcaend Mcaiacetiaadccia-eitn prptisreeclc.Adconomstubrae odadcuehinaeadmitraisvr sorption are low. Temperature resisiance of generalpurpose types is intermediate betweca that of polyesters and phenolics. Formulation anid fabrication oy swt psiltesaexrmly vraie
epoxies can be formulated for uses tuch an wet
or prcpregs for room-temnperature or elevatedtempei ature curing, and for fire-rctardancy. The choice of curing agent plays a maijor part in determining curing characteristics, temperature resistancc. chemical resistance, flexibility. etc. In addition, a variety of modifierF and fillers is available to provioc specific qualities. There are relatively few disadsantags-s with epoxy rcsins. However, becaube an~ine curing agents that are commonly used in room tempeature curing forlmulations mady ctiuse severe dci matitis, skin contact must be avoided. MIL-R-9300 contains requirt.ments for epox) lamninating resins, wvhile requirements for epoxy laminates are covered in MI[-P-25421. 2-4.1.2.3 Phenolics Phenolic resins are used primarily in (3RP appli-
an inrinrpnsive material
%xhprc
,
(
th heat. re-
sistance uri to 50001: and/or rionflanimability is required. Excellent elctirical propcrtics also arc obtaircti. Because water is produced and released in the curitig reaction, relatively high molding pressure is req4uired in order to prc~eni porosity in phenolic laminates. Preprcgs ticarls alk-ays are used. Mil.-R-9299 and NIIL-P-25515 cover the requiremncrts for Phenolic laminating resins and phenolic laminates, respectively. 2-4.1.3 Types of Reinforcement Gilass reinforceniem is av;,1lable in several basic form,,. anid in a %ide variety of specific construo'tions ssithirr these basic cztegories. Those formis coidnion1) j.scd in 6RP lamirates include skoven fabric, chropped fiber mat, and! i~onwoven continuous iapes or ro,*ing. Neaýrl) all of these art: dcris,,d fromn continuou% filamrents of 0.00023. 0.00028, or 0.0003h in nominal diamecter. Numecrous standard yarn C(:ii structio11s are available, with %ar)yrtg numbers (if
3
parallel Miaments per strand, strands per yard, and twists pc.- inch of the strands. Likewise. there is a multiplicity of fabrics woven from thc.ic yaifls that vary not only in type and amount of ) arns but also in thc type of weave. MIL-Y-l 140 is an excellent reference for definitions and rcquit emcnts for ti,a various yarns and woven fabrics, The type of reinforcement selctcird will depend upon the mechanical property requirements, part shape, and applicable fabrication technique as discussed in the paragraphs that follow, 2-4.1.3.1 Nonwoven. Continuous Filaments axium Thi renfoceentoffrs fom o mechniclpopeties bu ha miimumfabicaion possibilities due to the difficulty of placing and aligning the reinforcemuent in complex shapes. A big adv'antage where this type of construction is practicable is that the fibers can be oriented in proportion to the stress in anly given direction. Filament winding is the most widely used fabrication technique with nonwoven continuous filaments. This construction is covredmoecmpetey n pr.2-43.As 2-4.1.3.2 Woven Fabric This constru-ctwn proidus good mechaniv-ul properties and formability and is. theref'ore, the niost commonly used reinforcement in aircraft fabrica-
2.4.1.3.3 hopilped F"be The third commo3n form or rcinfoiroment is chopped Fiber mat, as defined by MIL-M-15617. Because thc Fibers arc short inno their orientatIon is completely random. this material is veycm formable. For the same reasons. atid also because or its high bulk - which limits the percentage of glas obtainable in a laminate - mechanical properties are lower than with roving or fabric. Continuous (swirl) strand mat is another veriation and is particularly useful for deep contours. in both types of mai. th-glass is held in place with a small amount of resin binder. Both types are available in weights ranging fromn 0.75 oz to 3 oz per ft1. corresponding to laminl. atd thicnfiesseno absou0.3in. trodue 0.00in.per naldMith rei~nfo 6icemincallso ishproduc ldin promrn-
h
1 A_4
adsetmoigcontewihrsitiss pound (SMC). Chopped fiber parts can be fabricated by the sprayup or Preform t-echniques described subsequiently.
4..
FaratoMeod
previously discussed, each of the common forms of glass reinforcement (roving, fabric, mat) can be niorchip el either dry or preimpreitnated withth ) laminatirip resin, which is cured or dried partially to a ' solid or tacky condition. The latter, called preprep, tion. When wetted with resin, the cloth has conarc advantageous in that they contain a controlled, siderable ability to stretch and conforfm to rather uniform, and readily measurable amount of resin. complcr. contours. Although intended specifically for They arc, therefore, easier to lay up, because wet laypolyester laminates, MIll-C-9084 fabric usually is up operations often arm messy and odorous. Prespecified for larniiat"s made with all resins. IKaquirepregs can be obtained with varying degree of tack so ments for eleven basic fabrics and six subtypes arc deas to suit the spwcific operation. And, because cornfined in the specification. Approximate thickness per~ pleie quality con irol tests can be mitde before the part is fabricated, the. problems of incorrect weigtding and ply ranyes from 0.003 in. for 112 fabric to 0.027 in. mixing of the resin system arc eliminated completely. for 184 fabric. (Still heavies fabrics, woven from :he( fife-, pre-pre" er In o~z rovings rather than yarns, arc avia~hbl Ii- i'll-----ses formulated with curing agents or catalysts thatI of up to 0.045 in. pcr ply-. these are covered in MILrequire heat to cure (generally 250*-3S0*F for at least C-i19663.) Most of these fabrics are balanced %%eavcs. I fir). Under heat, the resin melts initially, and then with nonoinally equal construction in the warp and fill directions: 181 fabric at 0.009 Mi.per ply is the coiiverts chemiically to a thermoset solid. Some pressure almost always is required in prepreg lamistandard balanced fabric upon which most test laminating in order to maintain good contact betwtien nate- and published properties data are based. Repireplies of reinforcement. This pressure results in reater cm, ing the extr,:me of unbalance is 143 fabric. %hich resin flow and, consequently, in higher glass ratios has a warp strungth about 10 times as great as its fill strength. This approaches the nonwo%,en conarid better mechanical properties than are obtained with unpressurizcd wet-layup laminates. struction described previously, sacrificing sonic mechanical properties for improved drapability. Generally, epoxy resins and preprcgs of roving. Most high-strength, glIAss- fabric- based laminates Lre tape, or fabric are associated with components of made from 181 and/or 143 abrics. higher quahlt, cost, and strength, while polyester As with MIL-C-9094, MIL-Y-l 140 originally swas resins and wet-lay up (or SMC) processing of fabric or ) interldeo lor polyester laniinates. However. its renina arc used where maximum required properties do S quliremeflts also usually are specified for fabrics conri)t justif*y the increased costs. Those fabrication taining epoxy compatible finishes. mecthods applicable to construction of laminated *.o
n'
.es-m-l
2-13 u
F
.
W, v
2
j
AMCP W16-202
r
TABILF 2-5 PROCESS COMPARISON ;Ii)1 F('R (;R! lANIINAl -:S*
,
-
, O
POLYESTER EPOXY
CONTACT •IOLOING __
_
VACUUM
.
.
_
BAG
"OPEN MOLD
._
I
__,-
SPRAYUP
_'_
_
POLYESIER
V•A1 FABRIC
30 45
MAT
40
T
I" L VI? R S•, E
RL.
70 I011
70101I10
)
60 10O 220 55 455ONLY 60 70 TO 2,0
POLYESER EPOXY
CONIlOU ROV'NG
I
0
_
I
EPOXY
CHOPPED
30
50 1000 1
70 TO 110
p•,,1
CLOSED MOLD MATCHED-DIE MOLDING S~EPOXY
MAT
FABRIC
PHENOLIC,
MAT
"POLYESTER,
FAPI:,CS, PRI'REG
EPOXY
LIMI ILI? DY SiZL AU I)CLM L Oi AUIOCLAVL 00IOL
100 10300
BYI _
_
3'i
, 4 1• .
••jo
Of
FRVSAEY.QN,1 FR'..V5AEEYHELIEISI TO 1CIt BOATIHULS_ t S-
1,.?..
MELAMINE, SI,1N1 EPOXY[
PHLNOLIC, ,ELAVINE, SILICONE,
__IOW
L'11
_
22510300
CE POA! HUHl •.
50
ER PREFORM I,• EPOXY •i-
10 ()AIL
P
55SKBM _
ROIIUUIS Si"L 01 I'0
0
5 ROLYESTER
FABRIC AOE N ROV ýOVE ,MA 1 MA FABRiK EPOVXY R00 POY WOVEN ROV
Epoxy
••POLYISI
.
j
TRr
40
OVEN ROY
EPOXYFARI__ POY V.OVEN RO\'
PRESSURE
HAND LAVUP BAG .,...AICAEPOLYESTER • -': AUTOCLAVE
:-
_
A
f -IBI RGLAS I'E RR[ F 1.r, WERGA R61 I
RESIN
PROCESS
•
R,
SP.
L1
'25 10 350
100 10 3000 1 ' 6 .ANE f't, NLI S
2Vr'TO 350
IN
O
_
1O
{ "1
ANE LS•'U 1C5
AND NONV"OVEN
_j _._
FROM OWENS CORNING FIBERGLAS CORPORATION TECHNICAL BULLETIN I-PL-1998-B
GRP helicopter components arc jiscu::scd subsequently. General guides to moldinj; processies. and resultin, laminate properties, are shown in Tablei 2-5 and 2-6. respectively. 241.4.1
Opcm Mold Hand Layup
simoothed onto the exposed surfac of the layup in order to provide a better finish. 2. Vacuum bag. A film (usually poly,,inyl alcohoIlI er nylon, is phlced over the surface of the part and
scahLd at the edges, or the entire mold is pk.ccd in a bag. A vacuum k then drawn. resuittlg in the applica.ion of atmospheric pressure to thr"lamfinate. I:ven This method consists simply of placing the rethis re~ativ,:'; low piessurc (15 psi) conside•Fabl) intquird number of plies of reinforcement and resin proves th- laminat,: quaiity bh reducing entrapped over a single mold surface, and rubbing or rolling out air and resil|-ricin areas. the air. Curing then is accomplished by one oi the 3. Pressure b:i. In this case a rubber film (o!'Ten "following processes: C t l T n i wconioured to 117c part shape) is placed ovz-i dit- l, tip 1I.Contact molding. The laminale is allowed to -and thŽ mold is scaied snth a pressure plate. Air or cure without the application of pressure. usually at stvam pressure ol up to about 100 psu then is applied room temperature. Heat can be applied to accelerate to thl. cavitN. the cure, but the contact process usually is employed 4. Autoclase. In this variation: of the pressuwc t'tg for large parts and/or short pruduction runs. and process, the ent~re assembly (mold, layup, rubber "~'.•,both heat and pressure 11ay be impracticabic. A stripfilm) is placed in a steam autocime 'nd cLured, norpabla film, such ab cellophane, sometimes is malhY at about 50-100 i'. 2-14
1 '
706-202
___________________________________AMCP
TABLF 2-6 (,FNI-IT.A1. PROPER} IFS OBTAINABLIA IN (&ASS RHFN1FOR( I) I'l.AS'I(*S* POLYESTER
EPOXY
PHENOLIC
j(-LASS fMM. FPL FCR11l
GLASS MATT
rROPF RiY
OH SHEET MOL iINi3
CLASS CLOl H
GLASS MAT
GLASS CLOTH
SP~lCI l'CfRAVI] Y
1.315 -10 2.30
1*012.0
1.810O ?.C
1.910O 2.0
1.70 TO 1.95
CLASS CON EN1,'. BY W H TENSUL STRUNCJ H,
251TO15
460T057
65 TO 70
41 TI 1
15000G 10 25,000
30,000T0 ?0,0C0 14,OOOTO 30,0%0 ?0,000TO 60,000 4,.000 To 60.000
COMPRESSIVE STR LNLI 11, 1
15.o000vO 50.000
?5,OOOTO 50,000 130,000!0 38,00T130,00010 70,030 17,000 TO 40,000
FLEXURAL SI R[ N67. H,
25,000 10140,000
146.0 00 7TO 90,000 20.0 0T1 26.000 70,30010 100,0 00 10,0 00T . 95,0001
i475
6.5X0,5 il. NO1ICHED HAR? ot -!1 in .O[ NOTCH)l VWA[ER ARSORP N ,i Z411i Lhil. HICKNESS, '
81
TO030
05
-
BURNING RAIL VOL AýITIiTT 50. RH AND 73" F,owm. 'cnOi
U1.1i10i.6 uU -
-
.Q
-.
3(10 T-3 15
-
-
10-
04
11
120 TO 180
60 10 120
fU
33010O 500 ____S--*--
.8x 1 0 3B
125 10 140
iIlU. _
____I___-__
O--1EXTHIGUIS11ING
'IS-
.;~U. U.UjI,
.:
-
3-N TO 3
lol
8 TO035
1 TO 26
8 TO 15 rrnnrt
o-n-
I
COI IE~ WNUOUS
ARC RESISTANCE,
O
330 TO 500
I-~..'-4 .XC _
5
___l"
100 TO 110
_
_
350 TO 500
-F-NONE XI 71
1
20 TO 150
'CC1MP!LFD 1Y RF INFORCED PLASTICS, COM~POSITES DIV.SOCIE3Y OF THE PLASiIV. IN('USTRY INC 2-4.1.4.2 Spi'nyup Ili this nethod. continuous roving is chopped into I1-to 2-In.lengths and blown, into a spraying Ntrearrn of resin and catalyst (hat is directed ugainst the mold (A fast roorn-tcrnilirature-setting polyestef generally Is used.) 1fhe mix~ture is land rolled to reduce air and icvec the surface. The resulting part is similar i11 con%.tructioii !o a chopped fib~er mnat hand lay up.
peratures to 350*0 F commnonly arc used. Prepreg. fixbvics and tapes usually are spr-ciied for aircraft applications requiring maxinium strength-to-weight ratios. I owcver, fo'r cornpkA) shapes and volume production. choppeci glass preforms (held together, like mat. b) a sinalI arniunt oý resin binder) frequently -ire used. 2-4.1.5 Surface Finishes
Althugh(hi C-1s i vey . uficentforlare cm-
-or man) applications, a smooth surface, free from aii pockiets and exposed glass fibers, is required. Such a surface may be specified in order to improve weathering characteristics, material-handling capabilities, human contact applications, or. 6riiplv, appearance. There are three different methods used to obtan a smooth, resin-rich laminate surface: 1. Veil mats. Thes.- consist of loose, nonwovf-n mats ot glass or synthetic fibers. THickness may range
poricots. it has limited applicastion in aircraft constructiooi due to pour Uniformity o! thickness and h-o-wightratos.outdoor lw relaivey s~ 2-4.1.41.3 %latched Die Molding
Whenever -'4)sely controlled thicknecsses are re-
quircd. two miold halves ar±. necessary Matched dit mrolding also is practical for high- volunic production even where high-quality surfaces and close tolerances are not required. Pressures of up to 300 psi and tern-
1W
70fi
inav be customn rnoldicd b) procedures sintilar to those dcsc:rihcd in par. 2-4.1 for low pressure, closeddie lamina.te, Industrial lamnizates are used for components of simple geomectry rcqUI.ing internmediate strength, lightweigh~t, and nonmetallic characteristics. In hell. copter construction, they frequently arc used for wear surfaces, such as oil conduits and lpuleIys for control cables anad iii electrical circuit boards. industrial laminates can be made with a number of' chemical, thermial, and electrical properties by varying the ty'pe and ratio of resins and reinforcemerits. Those combinations that presently are avsail~tblc coniinercially art; sho~ i in Table 2-7. In cach case, the laminates are manuifictured by stacking Lip. sheets of the irnpregnawed reinforcement (or by vrapping, in the cas;: of' itbes or rods) and curing tHicn undcr hecat arnd pressure. Vcry high pressure rainging from about 200) to 2500 psi used, resUlting wnhigh-quality, void-free parts. Fromi the biisic: moinbinations shown in Table 2-7 niore than 70 standard grajdes of laminates are derived. of these, 32 grade,, are classified by the
fToni 0.001 to 0.030 in. They are so loosely consiructed that resin content in the veil are.; is thbout 85 by weight. 2. Gel coati. This technique conlsists of spr'i,coating the mold surface with 0.0i0- to 0.020-in. layer of thixotropic (nonsaggiiig) resin, which is allowed to sct ayig rio uto theglas rinfocemnt. Resiiicint resins usually are used so as to provi~ie a compromise between scratch resistance ::nd impaL-t strength. Most gel coats are polyesters. bu! th,.methd an e usd aso wth poxes.mechanical. 3. Thermoplastic fiims. This method consistIs of laminating a film or sheet of weazhcr-rcsist::iit and/or decorative plastic, such as polyin)! fluoride or acrylic. to the GRP surface. This technique should be applicable to a vai jety of GRP processing methods with both polv-estci and epioxy rc:,ins. but it has not been used widely in the past. Re-cently. however, .i process involving vacuum-forniing of thermoplastic sheccis -which then are reinforced by %praying the bacck side with chopped glass and polyester resill has found wide acceptance. especially in the nianufacture of large parts (up to 300 ft-). FABRIC
-are
-A.MiNA US..
-
I
-....
A.cs
!at
Industrial laminates, also cýalled high-pressure laminates, are reinforced p~astics that are maimifactuied in standard. simple shape% such as sliccts. rods, and tubes. F-abrication of prits For materials generally is accomplf~ished h) standaird
(NI NI.\). Descriptioois of ihe 1NWMA grades and thenr applicatiois are contained in Vol. 46. Moderni
mtetalorking operations, sicias cutting. drilling.
po,- design. (icuicral characteristics lesiltinc' Iront the selection of the various reiniforcemeints and resins art: descrihed sibsequciltl).
punch~ing, and machining. In conrarst1
tO
'th
Lcdpi.asloaetepretesf
thcse laiminaites. The designer should consider grades. app~lication. aiid producihility prior to final corn-
Molding it)
the desired shape as discussed in par. 2-4 1- Where prooneicon quantities \5arrant inold costs., pairts aulso
I ABLF 2-7 F
RE SIN I YFF
REINFORCLE?'N1
PH tjOIL
FCR%*At D[[
Sh1.LI
Iý h
luf
HAI~YPE
C it-
I L0-I
I
CC)MTN FAWiC AS[BES 10 PAIq R
I
K:
~ ~
'1
I
0,1
S E. HE KO
iii. I
týcý
S:,K01,11 4-
C[
%tIL
I I TL112
I:OLlI
.
~
1~
-
ASBESTOS FAhkICI
NYLON FABFiC ..
1III
AND MAT
'FROM1 MiODRE N PL AST H Eii'C
L i zz
1
*
MATE RIALIS i-.S,'iPr., UJ.iLJ F L~[ ';I'. i.....
2-16
iIF
.II.'NS
NHI
-
AMCP 7W202
-
-.. ) .2-4.2.1
Reinforc4.ment Sdelltion
ParaLI'.el) low fur glass laminates, and c.ost is high.
laminates art paper, cotton. nylon, glass, and iisbcstos. Attribute%of these materials are: 1. Paper. I hie leas~t expensive, and adequate I-or many purposes. Kraft paper has rclatively long fibers and is thc strongest type. Alpha ciellulosc offers improvtd e~ectrical properties. machinability. arid uniforminiy, while rag paper laminates have the lowest water absorptiort and intermediate strength. 2. Cotton. Better impact arnd compressive .strengths than paper, and most grades are only slighily more costly than paper. Electrical characteristics. howevcr, generally are not as good. 1 he heavier fabrics have the be:st mechanicial properties, while the fine weaves have good miachinability. 3. Nylon. Low moisture absorption anid excellent impact strength and electrical properties, as well a,. good resistance to chemicals and abrasion. l-owevei. nylon laminates have relatively poor creep resistance at elevated temperatures and are coniparatively expenisive. 4. Glass. Highest mechanical .stre~igths by lar.
5. NKllainiiirc Excellent arc resistance at moderate cost. NMclhanicul priopertits, and heat, nlame, and -cheminia;l resibtancc qualities, also arc good.
Thcse materials also have superior electrical propertics~ and he-,m reskistnce. Cost is relatively hiigh. 5. Asbestosi. Used in the form of paper, gnaL, and
2-4.2.41 Speeifirstions lit addition to the NEMA Standards (Pub. No. I-I 1-19,65). thei foliossing M~ilitary and Fedcral Spccifi
~/ fabri-c. These laminates have excellent heat, flamne, chemical, and abrasion resistance. Costs range front low%foi those sith a paper base to high for a fabric base. The designer should select the type of miaterial best suited for the application. considering interface recIuirinlents mind felmt~bihty,. maintainmbility, pro'Jucibilitq, and surv.ivability.
2-4.2.2
Rvsir Setewicin
The rcsins ase6 in the manufacture of industrial ipihrdi.,. 1 Jui: c. .1-o!lyestcf. S!!(voome, and
-.
kRoomn temipcature nwcchaniCAl properties are con;-
The common reinfsoricenents weda in lrigh-pFC~Surc
meLAniinv. The characteristics of each arc: I. ieot he;~li idl~use Te b lr.ihe
resins are iriexptensise and have adequate reloperties (inccharotcal, electrical, therniul, and chenical) for inanN dcsian apiiin.irmlnrced 2.hpoxý tUsed csp"ejial v% here tcgh rcsistam~c to chemnicals and moisture is rcquired. Mechanical properties andi dinicn.,.ional stability also arc superior,. 3. 'olye~tei . Less coimion. but use~d 1icrnechainjCIIAJr1ld electrical a m.tn.espeeialls %%, e flame rcsist1tiiee is a1rcqu!(cnicirt. 4- SilcOiUC. Used p~rima1ri ld,wt glass fahm ic \Ahere iicelt resi!m-!icc ito S001~ i-- rc'jimrcd. Arc: esmmalirCc is1 v\Menlemt and 11(1i1roiv arbsiorpt ion is loss. 1 lie \c r lIs mA ivssipamt lol laciot ol the-se rc~smitv at higrh ItieqUefl~cs i, utiliWc rm radMi and r-±dio ins~ulator%.
:~
~~
2-4.2.3 Speckif Types In .ddition to the materials listed previously, there arc two special typq-s of industrial laminates that deserve mention: 1. Postforming grades. Made from resins that. altltough thermoset, will soften eteough at elevated temperatures to allow the material to be molded into intricate s!ýtpes. Sptcial paper or fabric reinforcenieit also is used, permitting considerable stretchilig
2. Clad laminates. Clad, on one or both surfaces. with a %ariety of materials, including aluminum, coppu r. stainless steel, silver, magnesium. and various rubbeis. "lhe corticr-clad laminates (gerierally glass/epoxy) are used widely as printed circuit boards.
cajions arc applicablc to fabric larninates.: I. ILP-S09, for sheets, rods, andi tubes of various resins arid reinforcements 2. MIL-P-79, for rods anu tubc-ý of paper/phenolie, cotton/phCF.olrC, and glass/melamiinc 3. L-P.5 13. for paper/phenoltc sheet 4. MIlt-P-I 5035. for cotton/'phenolic sheet S. MIL.-P-W824.
for cotton/phenolic sheet for
Aatcr- or greasc-lubricated beairings b. MIL.P-I5031. for glass/mclamnine sheetc 7. M: i -P-l15047, for nylon /phenolic sheet. 2-4.3
FILAMENT COMPOSITION
Thi., paragraph is concerned primarily with high-
pciforinance composites. consistinig of plastics rewith nonw-ovn filarnints of glass. boron. 11nd high-modulus graphite. Because the fibers areV itonisso-efl and, usually, kintwisted. they can be packed to high lither loatdings. The fibers (-an be Oriented .mloiig the axes of stress in p-rcrorticod to deSAgri requirements. allowing efficient ut iti~ation of the outistaniing properties of this type of rcinforcement. Whic~i the spccilic strength tiensile strengthv-to-densit) ratio) anid
specific modulus (Younig' modulus-
to-dcnsity ratio) of tire metal alloys (alumninum, steel. ialil
corumoril) used in aircraft construction are
cornilpared, it isýshow ii that they are necarly equal to 7~ to 9 X 10'!n. arid 100 to 110 X 10, in., respwcively.
I c
-,.
-
.1
W-Ak
Although S-glass offers a substantial improvement in botO quantities, niore important is the comparalively recent introduction of the exotic fibers. with boron and &;rapbitebeing of primary commercial interest. Unidirectional composites made from these fibers have specific moduli in the 600 to $00X 10'N in. range. Thc specific strength of boron composites is comparable to that of glass. while graphite cornposites are somewhat lower in this property.
2-4.3.1 Types of Relmorcemevit A summary of the properties of the previously mentioned filaments that are used in reinforced cornposites ~s contained in Table 2-8. The derivations and characteristics of these fibers arc discussed subsequantly. TABLE 2-9 TYPICAL VALUES OF PHYSICAL AND MECHANICAL CHARACTERISTICS O3F REINFORCEMENT FIBERS Lsw
MATERIA
Ius I -GLASS
0O92
S-CLASS
0.0oý
MEIA DEST
97--
t-LS OD9
'NTPGTN)0.093 GRAMlPTI0.6
GRPIE400 (PA~,
*into
EC11
Tj0J
5ID 50
-11
7.4 92
85
14 13
5-2,
50)
5.4
60.0
650
300
4.9
50.0
820
Ui±.
40.0..
-
--
fYE
.~. 635
MIHTS-904
24.3.1.1 F-glass This glass was developed originally for its superioc electrical properties. Glass roving is manufactured by drawing the molten glass thiough resist ancc-h.-ated platinum bashings at about 2400*F. From 51 to 408 (usually 204) filaments are gathered into a single strand, coat2d with a binder, and wound onto a drum at approximatcly 10,000 fprn. The coating bonds the filarents a strand, protects them from abrading each othcr, and also serves as a coupling agent to improve the resin-glass bond. For use with epoxy resins, an 801 sizing usually is specified. Requirements for Fglass roving are. contained in M IL-R-60346 under the Typc I classification. Standard continuous roving uses ECG 135 strands - where E designates the glass composition, C indicates continuouis filamenzrts, and G dcsignatcs a filamerit diameter of 0.00037 in. - resulting in 13,500 yd of strand per lb. ECG 67.5 (408 G filaments per strsnd) and ECK 37 (408 K filardunts of 0.00)052 in. 2-18
n
N
243.1.2 S-glass This composition, sometimes called S(994), was developed under Air Forct; contract for its highstrength properties. S-glass is available in the some forms (roving, tape, and prepeg) as is E-glass. The standard roin~g decsig'iation in this case is SCG 150,
niaigta there are 5.000 yd f strand per l
10 0'~vs
5.A
5s
diameter per strand) ravings also arc avuilabic. A roving package is made by winding a number of strands (or ends) under each tension onto a cylinder. The number of ends ranges from 8 to 120. with 60 being the most common quantity. Staniard packages range from 7 to 35 lb nominal wveight. E-glass rovings are available widely, both dry and preimpregnatec! with a variety of resin systems. Prepreg tapes of unidirectional filaments up to 48 in. wide, having a nominal cured thickness or -ither 0.0075 in. or 0.0 10 in., also arc available. These can also be purchased in tuo-ply bidirectionai (0 deg. 90 deg) or three-ply isotropic (-60 deg, 0 deg, +60 deg) forms.
due to the lower specific gravity of S-Slass- The major d0est46krrent to its wid,, use has been its cost. which is about i 5 oimtes that of E-siess. A cuumiiwiciu giadc.
containing most of the S-glass properties at
s.ew, lower cost, has been introduced. Another development is 970 S-glass. which has
20% greater moduius and ultimate strength than SThe chemical compositions of various glass reinforcements are presented in Table 2-9. IITS-901 and are the epoxy-compatible sizings for Sglass, while 470 sizing is used with S-12 rovings. Sglass rciving requircinecits also are contained in M ILR-60346 under the Type IlI classification.
2-4.3.1.3 boron Filaments These products currently arc made by vapor deposition of boron on very fine tungsten wire. Work is under way to develop boron filaments on glas-1 or graphite substrates in order to reduce cost and total density substantially. In order to make handling practicable, the material usually is supplied in tollimated prepreg tapes that are one filament tihick and up to 3 in. vvidc. A Military Specification on boron filamen~t preprcg is MIL-B-83.169. 2-4-3. 4.4 Craphlte A wide v'ailety of filamentary carbon p~roducts is produced by pyrolysis of organic fiber!. These prod4ucts may be divided into two broad categories: low-niodulus and high-modulus materials. Lo%% modulus carbon and graphite are iused freqjuently in
-
AMCP 706-202
NOMINAL COMPOSITIONOYCLIASS
TYPE
SiO., A12 03 MgO0
E-~GLASS
54.3
15.2
4.1
S AND S-2-GLASS
64.3
24.8
10.3
970-S-GLASS
-modulus
*
graphite crystals. Material developed in the U S usecs ra!vyQn fibers and hasq an irre~ular trpopcorn shape) 3~=. cros sctin. Matti ial developed in Engiand i, pyrolyzed from a polyacrylonitrile (PAN) precursor having a circular cross section. In either case, the average filament diameter is 0.0003 in. The Brilish PAN material is made in untwisted tows of 10,J00 filaments, and is available in continuctis lengths. The rayon-derived, high-modulus graphite used in the U S is made in continuous leng~ths from 2-ply yarns having 720 filaments per ply and 1.5 or 4 twists per in., depending upon th~e manufacturcr. The greatest development activity in highpromanc. flie s lfI'-,su~e uponi graphite. Dupimtn the small filaments, it can be formed around radii cts *sinall as 0.05 in., a major advantage over boron fiber. It also is expected that the greatest potential for cost reduction and product improvement lies with graphite. Evidence of both was displayed recently in the comi.:etrcial announcement of a 75 X 1IV psi modulus fiber at $400 per 1b and a 30 X 10, psi fiber at SCO per lb. Laboratory quantities of 100 x 106 psi modulus fiber have betri produced. Because new products and new manufacturers fre(luently enter the field, the data ini Table 2-8 include only those products with which a significant amount of experience exists. G(raphite fiber can be pr,-duced in the same variety of forms as can glass Fiber. Thus, in addition to yarn and to%, fabric, mat, and chapped fiber can be supplied. As with boron, however, the most practicable form for most applications is unidirectional prepreg
z
-
-
CiD
BD 3
7.
8.0
-
-
62.0 19.0 9.69.4 1--
woven forin, which is produced directly from rayon fabric at a fraction of the cost of high-modsulus. graphite. Hlowever, these products. used primarily for -high-tempercr-mure insulation and ablation. have no known applications in helicopter construction. *High-modulus graphite fibers are produc:ed in a three- or Foor-step heating process. During tme final step. graphitization, the fibers are held in tension, thereby imparting a high degree of orilentation to the
-
GeD
tape. A Mlilitarý Specification on high-modulus graphite fiber preprcg is MIL-G-83410. Among the most serious disadvantages of graphite coinposites are poor abrasion and impact resistance. 1 hiss. surface protection frequently is necessary. Allso impeding the exploitation of this material until -cr% recently has been its low interlaminar shear strozngth due to poor resin-fiber bonding. However, surface treatment% have been developed that result in %hear strengths above 10.000 psi.
2-4.3.2
R hiu, While all of the resins discussed in par. 2-4.1 have been used ir filament winding, epoxies are used almost exclusively for aircraft applications at normal operating temperatures. Where nonwovcn, high-performance reinforcement is ustd, the bcst available resin system also should be chosen since the difference in resin cost represents a very small pcrcentage of the total part cost. Phenolic and polyimide resins are used only where very-high-temiperature operation is specified.
24.3.3
z
Manufacturing Processes
Structures of nonwoven reinforced plastics ma) bc formed by filament winding, tape wrapping. automatic tape layup, or hand layup. Filament winding can he performed with glass rovings. gr;aphite yarns. and boron singlk filaments. This process is practicab-le for %cylindersand tanks with high hoop stresses. however, it is limited to hollow structures with convex surfaces. Normally, filament winding is accomplished by rotating the part on its axis as on a lathe. Parts also have been wou-id by revolving the spool of reinforcemeni around the fixed pail. (ieneral~y, prepreg is used, but wet winding also. is practiced. In the latter case, the reinforcement travels through a bath of high-viscosity (at ambient temnperature) resin system that is healed in order to lower the viscosity for efficient wetting of the reinforcement. When the impregnated reinforcement is coolet'
2-19
K-
AMCP 706-j22an
'...
- . .
to ambient temperature, the high viscosity is icattained. Latent curing agents must be used in order to obtain a rcasonablc pot-life for the heated resin sys-
various rcinforccnitnts in such p-roportionsan orienltations as arc required in order so obtain #Imost any intermediatte properties. The possible
tcm. T41K winding is similar to filament winding, cx-
effects of differenit thernini expansion coefficients must be considered, howvver.
inforcement (generally 1/8 in. wide) are wound. A recent advancice in fabrication technologyi unr ically controlled tape-laying machitie capable of applying prcpreg tape (hicated, if desired) at a controlled rate and pressure, and shearing it at the desired ls-ngth and angle. Still another machine applies reinforcement in three dimensions by weaving fibers perpendicular to the normal laminate. This, of' course, greatly increases interlaminar properties, which usually arc limited to the capabilities of the
construction of the spar envelope, skins. trailing edge, etc.. of rotor blades. Design studies have suggestcd the use of boron andhgmous g rpien these same areas, as well as in rotor hubs, swash-plates, drive scissors, transmission housings. drive shafts, airframe stiffeners, and entire fuselage sections. Boron hardware development presently is more advanced than .hat of graphite because the material was introduiced earlier. However, graphite coinposites are expected to be useful iii many of the same
resin,
applications.
Hand layup is still the miost widely used method where winding is not practicable. This process is no different from conventional layup of glass mat and fabric, except that the fibers are nonwoven and oriented, and generally are preionpregnated with resin. Filament-wound parts usually are cured under wrapping tension pressure only. allthough thicy mnay be autoclavedi or vacuum-bagged. Parts that are laid up (rather than wound) may be cured by an appropriate method as described in par. 2-4.1, such as pressure 'bag, vacuum bag, autoclave, or matched die molding.
Considerable design and physical property infortonicnaneinMLH)K?.PrI ndn Re.4
j
r4
2-4.4
HONEYCOMB AND SANDWICH CONSTRUCTION Sandwich constructio~n, ass shown in Fiog. I- is a composite structure comprising a combination of allternating. dissimilar, simple or composite materials, assembled and fixed in relation to each other so as to obtain a specific structural advantage. They are made of three or more laminations of widely dissimilar materials that can be considered homogeneous when bonded together. The layers include the facings, the bonding agent, and the core. The primary functions of'the core arc (1) to separate the outer layers so as to obtain a high bending stiffness, (2) to oupport these outer layers (the facings) in order to prevent elastic instability when they are highly stresised, and (3) to carry shear loads.
24.3.4 Applications Nonwoven, oriented filament composites are in order wherever maximum strength and/or stiffnessto-weight ratios in specific directions are desired, Thyar otuo~! ratc~~ wl-e nroyisrquired. Typical properties of thes components are shown in Table 2-10. 1 is entirely possible to mix the
TABL 2-1K TYPICAL UNIDIRECTIONAL COMPOSITE PFROPERTlE6 IJASFl) ON COMMERCIAL PREPRE~S
CFoLEXURAL
PLY.-c
FIBER IENSIL E 060511 IF1Xl :A[I0ltSSIVI COI4EN. I PLNG-l, MODULE,STRENCOI1, MOVLIU. STRLIIGIH to Ps ki %______ VOL k,si 1 i' V. ps i k,s 10ps O 61j 109 71. 200.0 7,0 99
4
L -GLAS.i
0.0015
S--GLA5S
0.0015
63.5
BORON
0.0052
50.0
220
0.1
186-232 21.6-3M.9
230.0
1.9
120
245.0
20)
443 460
82)0
SAV-I L SPICIF IC S0 1RENGTH. N ooiUS, OCIH IGI in106.q 6.01Z9 Z.54 I0
9.00
0.0122
s
bIn.r
117.020D 16
3.04 2.45-3.05
12o 390-420
HMG.90
0.0m0
51.01
120
25.5
116.0
25.n
00
7.95
0.054F
2.22
47i
t:
THORNEL -50
0.3011
53.0
104
292)
116.0
24.1-,
-
1.40
0.0536
1.93
465-
:,
NORGANITE-I
0.0130
43.0
-',
821
0.0104
-
-
0.0130
52.5~
0.0552
-
-
S2-'
--
MORCANiff-Il
12 -
___-___
163.5
"'FROM3UCOMPANY'S SCOTCHPLY TECH4NICAL0DATA SHEETS FOR EPOXY PREPREG$;I1009-26 2-20
010 SHEAR
1.2.96
RESIN ONGLASS ARDRESIN OORONQ 1AND0 GRAPHITE.
5.
AMP706-d20
-
A
BONDING MATERIAL HONEYCOMB CORE
2A
IA Fiue21
adwc
tutr
Proerl deignd sndwch onsruFigurhas.mandwinch Stue atreat1tiehavr Rf5.
advantages; high strength-to-weight dnd stiffness-toweight iatios are the most predominant. Secondary advantages include fatigue resi ýiancc, impact resistance, and aerodynamic tfacizricy. A comparison of minimu-I-wvight design for
~ IA.
Honeycomb sandwich is the lighitcst possible material that carn be used to achieve an optimum stiffness-to-weight ratio. A com~parison of various materials (Fig. 2-2), based upon an equivalent deflection. suggests a 30% weight advantage when comn2-21
'
AMCP 706-202
3600
~limited,
oneycom b core const i uct ion represents by far the
______________________I _________
MATERAL j HONEYCOMB SAND~IICH 0.058 NESTD 1 EAMS0.05 NESTD " ' EANS0.08 STEEL ANGLES 0.058 MAGNESIUM PLATE 0.058 ALUMIUM PATE0058 STEEL PLATE 0.058 GLASS REINFORCED 0.058 PLASTIC LAIAE
mnost cases, no more than 3/16 in. Among the materials available for sandm~ich zipplication are conventional honeycomb. foamis, and balsa wood. Advantages of balsa wood and f~m are and their use usuaily is due essentially it) a limited physical characteristic requirement rather than to an overall property consideration. Balsa wood is used predominantly in flooring applications, where the need for continuous support is provided by the fibers. Employment of foam cores in a sandwicht construction is.essentially, a cost consideration. Both b -alsa wood and foami may produce adverse effects. and also may limit the environmental capahilities of' the construction.
WEGT1b 7.79 10.6 1.86 25.90 26.00 420 68.60 83.40
Figure 2-2. Weight Comparison of Materials for Equal~~Doinsuch advantage when compared with a flat aluminum plate.Fiberglas Optimum fatigue resistance is a byproduct of sandwich application. The increase in flexural and shear rigidities of the construction, at no increase in mass. provides for an increase in the fundamental modes of ecttotohigher octaves. In addition, the attach-
most efficient utilization of parent material. Conentional honeycomb cores, as illustrated in Fig. 2-4. are cs-witially hexagonal in shape and are manufctured from almost any material that can be made into a foil thickness. Properties of' hun~ycomb cores can he predicted accuratldy, based upon the con. figuration and the parent material properties. The erits of one type over another are related to the properties of the foil material; the relative increase in efficiency is related directly to the increase in the prpryo1 h oe i one~ycombh core material can be made from metals such stainless steel,ihrsn and titnu rfoas aluminum, iega mrgae --
as nylon-phenolic and polyimfide. Other types of core material include those made from Kraft paper and Dupont's Nomex* nylon-fiber-treated materials. core material provides radar transprncadatssadilti.Ithsowiecrc cntnsadalwls agn.Katpprcr cosatanitlwostngt.KftPrroe mtra saalbei ayvreis n sue 200
uAl21
ment of the core at the facing provides visco-elastic
3mi
damping that prevents amplification of resonance. it ca180 is quite possible to design a sandwich structure for an 3 .1 . AT ifintRe life under cycling loads, provided that the 163 (HALAC maximum loading is no more than 35% of the ultiU~ 160\zmate capability of the construction. A comparison of ~ 20 hr AT conventional sandwich structures (Fig. 2-3) in a sonic environment indicates that the sandwich can operate
*
in excess of 500 hr at approximately 160 dB, (decibel) while skin-stiffened structure of the same weight will fail at less than 200 hr under 130 dB. Aerodynamic efficiency of a sandwich structure is a consequence of the continuous, uniform support of the core materials. This characteristic is accepted widely in both aerospace and aircraft applications. Vertical support of the material in the sandwich construction is limited in span to the cell size, which is,in 2-22
-65
,3(*VADTI
T20 -
F I'VD
It -
TP%-5
-
TUCUl -
162 (1
~%%4
____
C) 140.
460 hr AT 167 OB ___
SRCTR SKIN-STFFENED_______R
120 0
100
200 300 400 -500 TIMIE, lir Fiue23CoprtvSncFage Fgr2-.CmatieSncaigeResistance of Conventional and Sandwich Structures *Registered Trademark
..
AMCP 706-202 when cost is u factor and/or thermal conductivity is of concern. Dupont Nornex nylon-fihbertreatcd core .matcri;.l, though recently developed, has thermal resistance and the properties required for aircraft flooring applications, Employment of a honeycomb core material in a construction is an exact technique. Physical characteristics of the construction must be investigated thoroughly, and rclatcd to available core properties, piior to the firming oi the design. In addition to the structural requiremen;s, the environmental operaring conditions must be explored, Common honeycomb types (Fig. 2-4) include the conventional hexagonal shape, a rectangular flexible core, and the reinforced and square cell shapes. The rectangular core is, essentially, an over-expanded hexagonal core. The flexible core is a configuration departure in that it inctL"Jcs a free sine wave that allows the core material to assume compound curvature at no sacrifice in the mechanical propcrties of the
an additional flat sheet in the center uf the hexagonal cell so a%to favor a mechanical advantage in a specitic orthotropic direction. The square ce:ll core is a consequence of manufacturing case, and is employed primarily where resistance wclding techniques. are rcquired in order zo develop the core material. Although a predominant use of honeycomb core material is for constant thicknesses (flat, single and compound curvature applications), it also is used for such components as airfoil sections. The mechanical properties of the core material in a sandwich construction also must be considered. The core, whether isotropic or orthotropic, may be considered as a continuum spacer for the membranes (the facings). Typical properties of balsa wood cores arc presented in Figs. 2-5 and 2-6. Figs. 2-7, 2-8, and 2-9 illustrate typical properties of hexagonal aluminum core material. Several different alloys are presented, Table 2-11 is a presentation of the propcrties of typical rigid foams.
foil material. Flexible core, unlike anticlastic hexagonal core. does exhibit characteristics of a syn-
The term sandwich construction describes the close attachment between face and core material in this
elastic material. Reinforced hmxagonal core employs
type of structure. Should this attachment be weak, or
RECTANGULAR
I
r
HEXAGON
FLEX-CORE
SQUARE
REINFORCED HEXAGON Fi,
2.4. Commoa Honeycomb Configurations
"VP
2-23
AMCP 706-202 absent, the construction is no longer a saldwich. Attachmcnt of the core to the facings is necessary, and must bc of sufficient strength to develop the full mechanical properties of the sandwich construction. For example, if the construction is loaded to its limit, then failure is expected to appear either in the facings,
or in the core, or simultaneously in both. However, it cannot appear in the attachment between core and facings. It is most importdnt for the designer to investigate the properties of the bonding agent so as to assure compliance with these requirements. Adhesives of various types and propelties currently are available to satisfy every sandwich requirement. Table 2-12 contains a partial listing of
•.2500
common adhesives currently in use. Laboratory shear BALSA WOOD
II
-
- 2000
I
t 1-
,presented
, 0 Imaterial
:000 ,,
S500
2
o
4
6
8
bond strengths at room temperature of aluminum-toaluminum bonds with varicus types of adhesives arc in Table 2-13. Useful temperature range and strength properties of structural adhesives after exposure are listed in Table 2-14. process of applying adhesive to facing or core must not be ignored. For the adhesive to be
•
-The
10
efficient, it must be applied to joining surfaces that are free from oxides and contaminants, and its appli-
14 1G
12
DENSITY, lb ft3
cation must take place under controiied en-
Figure 2-5. Properties of Balsa Wood - Compresic
800 8
Strength ,s Density
-
I-
600600 .ALSA
WOOD
400400
-1-
•=o• 00
_ _
0
(,
5056 AND 2024
4
2
6
20
0 .-
•
8 10
12
14
16
52
/200
-"
DENSITY, lb It'
-
0 Figure 2-6. Properties of Balsa Wood -- "1 Strength vs DensitFl
2
4
6
8
HONEYCOMB DENSITY,lb.It 3
Shear
10
Figure 2-8. Typical "L" Shear Strength
=-2000 1IGO
-0
S150o
''-
-
10.,
Co 80
U 500 C
,I.ib52, 40
L
505b. 2024
40
0 I,.-HONEYCOMB
2
4
6
8
10
DENS!IY, 1b It'
Figure 2-7. Typical Stabilized (omprev%.ie Strength 2-24
U
2
4
6
8
HONEYCOMB DENSITY, 1b f3
Figure 2-9. Typical "U,"Shear Modulu%
10
1cM
OFRIGID OAMS TABLE 2-11. PROPERTIES OF RIGID FOAMS'.
DENSITY
DENSITY,
COMPRESSIVE STRENGTH.
SHEAR STRENGTH
Ib/ft3
psi
psi
MIN
MAX
MIN
MAX
-
CO 2 BLOWN URETHANE
1.0-I.1 65.0
FREON BLOWN FR(,, t1.5
M1N
MAX
POLYSTYRENE MOLDED
2,100
0.21
1.00
600
450
152 4.0
1525 15D0
1,500 100
10.20 11.00
200 65
0.11 0.11
0.37 0.16
350 350
250 250
1.3
4.5
10.0
140
15.0
95
024
0.33
115,
0.5
10.0
8.0
200
13.0
90
0.77
175
46. 4510.0
120
0.24
38.0
5
6,000
0.65
15.0
25.0
600
3,000
__ _-___ S
'L
IC
ON
_--__ E
_
_
.. .0.'11 2.0
8.0
13
110
I10
251
LHEAT E ACTIVATED POWDER
12.0
. ROOM EMP LIQUID
3.0
PHENOLIC.
LOW DENSITY MEDIUM DENSITY
.. 1
.
.....
2 .0_
HIGH DENSfTY
"vironmental conditions. The elapsed time betwcen "preparatory cleaning for bonding and the appli-
'"
cation of adhesive must be held to a minimum. Procss control during application and throughout the bonding of the construction is vital for the devciopmcnt o" the spcciificd properiy for the sandwich Adhesive manufacturer recommendations must be adhered to methodically. Design considerations for sandwich structural compon nts are somewhat similar to those for homogencous material. The main difference is the inclusion of the effects of the core material. The basic design concept requires the spacing of strong. thin facings far apart in order to achieve a high stiffnessto-wcight ratio. The lightweight core material having this propekty also will provide the required resistance to shear and the strength to stabilize the facings to their required configuration. Sandwich is analogous to an I-beam, the flanges carry direct compression and tension loads in a similar manner as do the facings of the sandwich, and the web carries the shear loads as does the core material. The departure fiom typical procedures for sandwich structural elcments is the inclusion of effects for shear propertie:ý "-on deflection. buckling, and stress. Because the
S.
-5" 40 o
0
024 j
.80
.60
0..ii 540 .850
35.. ... 360
650
-; , 2
14.0
I)
0.20
250
10
30.0
020
250
1,100
90.0
030
200
.
.
.
..
115
'.v0
.
.
-
500
0.30
325 -
.LUm3 ur.ru U II.,,,j
0 -~o... 360 38.0
I
..
10.15
PRE-FORMED PACK-IN -PLACE
MAX-
18,000
SELF E".PANDED
EPOXY
M
10.20
1.5
EXTRUDED
.
-
hr-ft (Ff/In.), SHORT I FULL MIN MAX TERM TIME
-
"
MAX TEMP,OF THtRMAL CONOUCTIVITY
022
..
I.30 '
-.
facings arc used to carry loads in a sandwich, prevention of local failure under edgewise, direct, or fiatwisc bending loads is as nccessaiy as is prevention of local crippling of stringers in the design of sheetstringer construction Struiurai instabiiitiy of a sandwich construction can manifest itself in a number of different modes. Various possibiliti ,Iustrated in Fig. 2-10. Intcrccllular buckling (face dimpling) is a localized mode of instability that occurs when the facings are very thin and the cell size is relatively large. This effect can cause failure by propagating across adjacent cells, thus inducing face wrinkling. Face wrinkling is a localized mode of instability that exhibits itself in the form of short wave length in the facing, it is not confined to individual cells of Lcllular t)pc cores, and is associated with a transverse straining of the core material. A final failure from wrinkling usually will result either from crushing of thc core, tensile rupture of the core, or tensile rupture of the core-to-facing bond. If proper care is cxcicised in selction of the adhesive %ystem. the tensile bond strength will exceed both tfl tensile and compressive strengths of the core failure. Shear crimping often is referred to as a local mode
2-25
"' '
.
"
TABLE X 12. COMMON ADHESIVES IN CURRENT USE AOIISIVETYPETYPICAL AHSVYL TRADE DESIG144TION Ar 30
NITRILL PHENOLIC
EPOXY PHE NOLtC
ADHSIE TPE AHSVTYE
3MCOMPANY
INARMCO
METLBON0D402
ViNYL PIIENILIC
TABLE 2-14. USEFUL TEMPERIATURE RANGE AND STRENGTH PROPERTIES OF SYSTEMS MANUFACItIRERO___ STRUCTUJRAL ADHESIVES'%
FM47
AMERICAN CYANAMID
At 31 MELOLONO 105
3MCOMPANY NARUCO
AEROBOND 422
ADHCSIVE ENGINEERING
HT 424 I HYLOC 422
IAGE*
VIY
ME MELBONfl I 5471
NARMCO
IAF iG 3M COMPANY FM123 AMERICAN CYANAMID HYSOL HYSO L 9601 PLASTILOCK 717 BF GOODRICH IRELIAL1OND 711&393 1 RELIABLE mrc IHP 103 HEXCLL
MODIFIED EPOXY 250CUR
METLBOND 328 MODI~iEA EPx F 120.ALL 31.0SCA.,.
GOOO TOEXCELLENTr
30-1 7W0 6
00IO
FAIR TO GOME
PHIENOLIC
225
PHENOLIC
400
2004~9W
-67
130-W00
~
0 UNUOOIFIEO
HECL
HYSOL
PLEL STRENGTH
________
AMERICAN CYANAkIID HYSOL
HýYLO3C 901 B-3
TYPICAL VALUES LAP__________
PHENOICRLE
HP36IXE UNMOO'FIED kPOXY
USEFUL TEMP
PHENOLIC
$00
1___0__1____
13-00
POOR TOMEDIUM POOR TOMtDIUW
m0m
_____
PEOI
,-
____
MODIF IEDEPOXY 250
1540-M
20C'E MOIIEOPOYD 3W0 CURE -
InIl 0W30 5020
IEPOXY POLYAMIEJE
IO 306
GOO
________
GOO33 O_________ 20-_____31_
-OIMD
NARMCO
POO
______
12UREPSE
3MCOMPANY
CURL
t;*otfl~:~.-.~.-.
Lu
i
.......
instability for which the buckic wave length is ver short duc to a low transverse shear modulus or the EC?216 3MCOMPANY core.. The phenomenon of shear crimping o1curs EPOXY I'CLYAMIOF. HP 31( HEXCEI. quite suddenly, and usually causes the core to fail in I CO~ 81y5O ~shear. General instabihiy for configurations having POLYAMICEFno34 AMERICAN CYANMtDfnoeninCELxcept at the boundaries ~951 involves overall bending of the composite wall HPI~.1 301P3132AHECE coupled with transverse shear deformations. Whereas MOOFIIUREMAESFM3 AMERICAt; CYANAMID phenomena. general instability is of a widespread COR SPIRNG AT 41320 31.CEOLPANYutrellrbcln n rnln r oai RELIABOND 398-420 RELIABLE MFG 41 04IEXCEL
(
-
CORESPIICIG A 320 *.Ki
~AnHFSIVFS
TABL 90 SHER BND TABEN .18OSHEAR
A
3MCOMANYnature.
REtLARONO 310B
BOD M
RELIABLE MFC
-13. C H OF DH1I3E
NT
FlDEIE
ADHESIVE TYPED STHEART ADHEIVETYPE STRNGTH NITRIE PHEpLIC*
sufficient core zhcar riidity. The basic design p-rinciples of sandwich construe-
ion can be summart'Aed as follows: I. Sandwich facingi s/mal be at least thick enough
towithstand design streins under design loads.
VINYL PIIENOLIC
4200
EPOXY PHENOLIC
34030
will not occur under dsign loads.
UNMOIFID EPXY MODIFIED0 EPOXY-250 CURE
4500
MODIIEDEPOY -30 CRE 300of EPOXY~~~ 50o ____________of
POLYIMIDE
3300
*AVERAGE VALUES AT ROOM TEMPERATURE.
* ITEST
SPECIMENS ALUMINUM TO ALUMINUM, LAP JOINTS.
2-26
mknsofyi-
2. The core shell be thick enough and have sufflcient shear rigidity and strength so that overall sandwich buckling, exo~ive deflecion, and shear failure
______350
*
Premature general buckling normally is ,ucjbanuTiraiuwn
3. The core shal have a high enough modulus of elasticity, and the sandwich great enough flAtwine tensile and compressive strengtths, so that wrinkling either facing will not occur under design loads. 4.LAIE For cellular honeycomb corms where dimpling the facings isnot permissible, the cell size spacing shWl be small enough so that dimpling of either wall irto the core spaces will not occur under design loads.
In~ addition, selection of materials, methods of sandwich assembly, and material property used for
~
I4
AMCP 706-202 design shall bc compatible with the expected environment where the sandwi.h is to be used. For example. facing-to-corc attachment ihall have sufficient flatwisc tensile and shear strength to develop the required sandwich strength in the expected environment. Included as cnvironnlert arc cllct%of temperature, %ater and moisture, corrosive a!-mosphere and fluids, fatilue. creep. and any condition that may affect material properties. Additional characteristics - such as themmal conductivity, dimensional stability, and dcctrical continuity of sandwich material -- should be considered in arriving at an effective design for the intended task.
2-4.5
ARMOR MATF:RIAI•S
I herc arc available a variet) of armor matcrial, and matncrial combinations that can be used for pas,ive protection of helicopter%. Armor types, vsith appropriate Military Specification references and a relative comparison of cost, availability, machinability. weldability. formability, and multiple-hit capabilit). are summarized in Table 2-15. Table 2-15 also lists the areal densities, and provides a comparison of strength. hardness, shock and vibration, and resistance to corrosion for the various typce of armor
"
FACING FACING
HONEYCOMB CORE
CORE
t t tt
ttt (B)SHEAR CRIMPING
(A)GENERAL BUCKLING
4444
444
(E) CORE CRUSHING
(C) FACE DIMPLING i
D (D)
!
-
,--
SEPARATION ,-
ttt t
FROM
CORE
t t
Figu'e 2-10. Modes of Failure of Sandwich Composite Under Edgewie Loads 2-27
na
-
- .
___.
I
X I
II
-
.
..
I2QQ
II~g
~~~~~~~2xI2
.
.. . .. ..-
Q0C
Q,
.
2
2
2
4
~
0
X I
0"
-12
Y
I
00
~
I
IT
2
1
-
'R
___
____m__7m
001'1 0 0
Ii
a
z2 z
42z
I Z- Z
00
0
rI
ý E
z~ Z~
0
0
8..~ _____
.8
z
2
0 8
Iz 02
=1
D
&0 FF4
U
3~~~24 5.IE
F-
C.
~
4. .2~ .~
-w.)
jS-
0D
0,
Sh:
2-28.
NOLV--Jd
SI~
VN
13X
AMCP 706-202 TABLE 2-16. FABRICATION DATA FOR LIGHTWEiGHT ARMOR MATERIALS
0250THICKPLATESIZE
4MAY
335
K-T SILICC.T4 CARBIDE PLAST IC COMPOSITE
PLASTIC COMaPOSIht
PLASTIC CMPUS'TE 1AONOITEIiC TILE 28,,. DIAG5X10l
UONOLITHIC TIEL MONWLTMIC TILE l.T~. 4XWIPANELS EA6,,,.. 7 1AiR..
HAVE BEENMADE
PANEL
I HICK
1
VERYSIVALLRADIUS IN 130THDIREC'IONS INANNEALED CONO
Im. DEEPDISHES HAVLBU~N ExlvPLOSIVELYPOIRMI P
I
NC( NODEXPERIF
O AVAILABLE
-
AI
SHAPES MAY CURVED
SLIGHTIMPROVE-
EON-,
TC~LREENRE
FLATPANELS
TOGI5CEUTIRED
7TRCHCU17TING
PRONELDING, CEDRE REUIED
STAINAUSTENITIC LES'STEEL IMIC OR SU6%ERG(C ARC) OR LOA PERRIIIC HVTTROGýIE
GAS ORHPLATIA ARC_________
FRCTIGTECHNIQUES
ANNEALFO CONDITIONt HARD[EX HEAT TREATABL IELECTRCTE IMC GIVESBALLISTIC JCINTHEAT -TREATED
.- DIAMDO I DOLS~
BE NELDEB M
HHOSCAN'T
COCNEITION -STAINLESS NITH 700EPREHEAT
FJCTRODDES V
WEQIR 1', ECxCARBID10E i9ITHSPECIALCARTIISHTORCHME AREINC TAIED1 OQ IHIAI AID O2HHA FR HEAI TIP.,OPOT LEUSENJ MAY
FDRILLING
.RI L:AGf'OSSILE
USINGMASCINARY Iy 8s yjOT
WITHSPECIAL CERAMICDRILLS
YES
f-
RE'WINAIN
IISERTS THREADED
VARIOUS ýYPES or
FOI 7A~WN 3TRiCTLIAE
, IN FIBERZAO TYPE PER-i~-IPHEKALSUPPORTIINSERTS
BRAKETETRY A4iDTHREAODE IN8tIIINS
ITO
FOR ATACHIENTTHROUHBDLINGFIANGE -
IN BOLT71 COPESIok COWICI.-EEATIGIIIS
-
NOTIMPORTAILT I
-~n
TLS
-
0ELVINS,DP I""HIIIECYANICAL JOINTS
PAWE. JOINWKý NITmOES
-
.ýM
ALEtIT C1ONii
Z
mawlAu~s. Tabk 2-16 su mrzsfbiaindt tht a
be, cosdm
o
ror instalia'iion on heli.
Aotes
2-4.5.1 AvILk14 Meakuil Maw&t that cana N. cunsadereG for Ue dn armor design. aimi t~.cif iropertics. includc 1. Aluminum alloy. ProdlicrA ku splash from bulk.t impacts thar., other armor materials in comnn~ vv, is excx:%*vnaiy effectivc ngiainw~ yawed ard higto-obliquity impwcU. and is nonmagnelic. 2. Titanium. Nonnaagsscic arid ralaistant to sea w'a~er corrost*n. 3. Homogvcou' steel. Rotoid fromn a starl alloy with the toughnes and pucct of elongation naxsswy xo achwor a good rzintence to bodth punc-
Must BE 7F PRAMIC1 RELIEVEDor CON PR[55IQFN5UNDER BOILTý ME EAD WITH THROUGH CRMCI 130.TING
CONTINUOUS Pj ASTIC B'.1KING.PANELS WAY RE APPLIED T'1omo A PFRAMEUSING METALDOUBLER
CAZ WE OfACHIEVEDC Pv RELOIN'.WITH CAh' OE A-HIE. L, B~t EfLOIN,. U141ES3 IAR[Lh EtECTA3CE IN IIRV EtDING UNLESS
I'
PANELS UPTO02.-..
2TGi.n.SMALL OUVATURE XT;AES8T ALED INGON :.C THICKNESS END IN TN ANNE APO CON,I ,ION
IATIKES
EFE ROL OPCLSATLL
2fx6x%,...UPTO 0001I6
9,00"H..
P01,5
EXTrNTO~r.OMuu4 6AUE
TAL6,HAR, NrSS UORTS1CL
fIMAX
1
I1I's01HICK
,!) RDUS Fra'P CAL 30 APIIIIEA1T
1
ETTREA"E, TE
qIGii-I-ARDINIES SSTEEL
ueadsaln.I
ONE
TO BACKRIG MATERIAL SIZEAND Of REOUIRED SHAPE
WIHVARYING DiGREES VKA CE
sdtsignatedmantc
ale" austeilk aftoi am not mwAgnt. If thes steels are coldworked, bowevu, tbcy become mnagfatic. 5. Hard faced steel armor pl,;te. Comnposed of a h'ird surface overlaying a softer IMck~ing matcrial of tougher steel. It is somewhat moire effective. on a weight basis. against solid shot than isface-harwkwA armor plate, and can be fabricated, by spoial tithniques, to a curve. It is magnei. 6. BalI6tic nylon. Providaes czoluen protectiot. from fragnaesols and tum~bled projuctika. Balistic nylon pads or quilts can be owasýJre for replaemont of insulation and soundi-attenuatirg b~les W e~ad/rtaIcn The ballisti level of nylo 2-29
AMCP 706-202 with fastcnc~s and/or attachments should be etablished and/or verified by gunfire tests for each confguraion.at 7. Ceramic. Built up of various materials, each intended to perform a particular furiclion in defeating the projectile. for example, a glass-fiber-rcinforced plastic to absorb the energy of impact, faced with a layer of ceramic tie (aluminum oxiide, A 1,0,; silicon carbide, SiC; boron carbide. B4C. titanium dliboride. TiB., etc.) to shatter the projectile. On a weight basis, some of these composite, compare favorably with standard steel armor plate for stopping solid shot. However. they have poor capabilities for stopping mnultiple hits, and produce many secondary fragments when struck. Ceramic is the bulkiest of the matcrials listed here. and usually is the most expensive. S. Ceramic-faced. The ceramic facing may be applied before or after the armor metal has been shaped or formed. 9. Transparenit. Composed of glass or clear organic polymers. either alone or in combinal ion (M IL6-5485. MIL-A-7168. MIL-A-46108). In general, ceramic armor exhibits the lowest weigpht per unit area for protection against armorpiercin~g ammunition 'cal .30 and .59)). Metallic armior exhibits substantially better multihit capability, although the probability of a small panel of aircraft armoi taking a multiple bullet hit from a high-firingrate Sun is remote. Metaflic armor for aircrew seats may become~ competitive on a weight basis when the arirror is used simultaneously as support or structure. figUred
*
2-4.5.2 Deslga For desgiv strcnith and rigidity requirements, refer to NIIL-A-88&0 and AMCP 706-170.
2-5
ADHESIVES AND SEA LANTS
2-5.1 BONDING AGENTS There are literally hundieds of proprietary adhesivc formulations suitable for various aircraft bonding applicationps. Some of these may be used in bonding a wide variety of materials, while ot~crs are usable only for highly specialized purposes. Adhesives generally are categorized under the two broad classificaions of structural and nonstructural types. 2-5.1.1 Structural A&Wnves Thi, category of materials is used for bonding primary structures that arc subject to iarge loads.'rypical ultimate band shear streirgths are several thousand psi. Structural adhesives usually arc formulatc~d fromt thermosetting resins that, when mixed with a suitable curing agent, react to form an infusible and 2-30
insoluble solid. Depending upon the type of curing agent, the conversion may occur within a few minutes roomi temperature, or. at the othcr extreme, it may reqluire heating up to about 3500 F to effect a cure within a reasonable time. The latter type of material, due to its low reactivity at low temperatures, can be premnixed and stored (often, under refrigeration) as a one-comnponent system until used. Nearly all applications of structural adhesives requirc lixturing in order to hold the components being bonded in contact during cure. this is because at some point during the cure cycle the adhesive goes through ai fluid flow stage. Most structural bonds arc made with tape or film adhesives. These are usually from 0.005 to 0.015 in. thick, and may be unsupported or supported on thin, open-weave fabrics (carricirs) of glass. !,yon. or other fibers. Filmn adhesives have two important advantages: 1. Uniformity. Variations .n thickness and cornposition are minimal. Because both shea, and peel strengths are sensitive to bondline thickness, control of this variable is desirable. Although bondline thick-1, ness olso is affected by curing pressure variations, film adhesives. - particulzrly those having flow restricted by carriers and/or high-melt viscosities can reduce thickness variations appreciably. In addition, film adhesives eliminate the weighing errors and inadequate mixing that are possible with twopart liquid adhesives. Quality control checks can be mnade on each roll of film before production parts are bonded; this is not practicable to perform on each batch of most liquid adhosives due to limited potlives. 2. Ease of assembly. Film adhesives are available in a wide -ange of tacks. varying from dry to very sticky. Complicated parts can be assembled simply by cutting the film to the shape of the desired bondline and laying it on the first suwface. The second surface theni is placed in position, and is herld by the adhesive tack until bonding pressure can be applied. Films that are not tacky at room temperature are tacked readily by momentary contact with a hot iron at sirategic locations. Parts of many layers may be laid up in this manner and bonded at one time. Adhesive waste also is mi~nimnized when films are used because there is no exmrss material left to set up in the mixing container. Most film adhesives require curing temperature. of 250*-3509F. and. thercircr. have long shelf lives. Cold storage usually is advised, however, although some types are stable for many weeks at room tempetature. A few types are available that curm at lower itemperatuics, including room temperature; these must be stored at temperatures well below 0*F.
I~
(
CtMCP 706-202 The other common phy stal form for structural adhesives is the two-part liquid niixturn. These mnatrials consist of two componcats that icact. when mixed, to form a them mosetting solid. Us~ually. tlhey are 10017 nonvolatile. Many cure at roomn ticmpe.'ature in a few hoars or days: other-, requir' hcat to curc Frequently, they atre in the form of high-viscosity pastes containing inert fillers und/o; thixotrofir. agents. In contrast te most film adhesivi.s. however. 'iesc: uncurcd pastes usually becomie fluid whet, heated. Pot-lives, like cure ti.edepend upon the rate of chemical reactivity, which is influenced greatly by temperature. Thus, adhesives that cure rapidly at room temperature may isave only a few minutes of pot-lf.fe while those requiring hightemperature cures have pot-lives varying from hours to months. Less common structural adhesive forms i-iclude onc-component pasle:; and powders, and all icquire elevated-temperature cures. Essentially all structural adhesives of interest for hczlicopir~r applications are based upon either epoxy or phenolic thermosetting resins. Because these rn~utpritik ire bruitt,. ~inhernith thri' iav;i~ly aire-,i~dt with elastortiers or thermoplastic resins in order to improve peel strengths. Polyuretharie adhesives also show promise, as they can be forinulated with both stre-ngth and flexibility. To date, however, they have not been used widely ini structural aircraft applications, Epoxies atz- the most versatile and widely used structural adheaiives. They have excellent adhe.:ion. low creep, low shrinkage during cure, and 100%~ nonvolatility. The liquid or paste typres have either low peel properties or '.empc.-t-urc resistance, and
'~fiexibic
J
arc
gvb
iadi*PAoto.
ritM,ificmuasautnio
(t`01 ainnprovcmvs-z
of these particclar characteriso'Ths) than are the film types. The latter can be modified with tough thermoplastics such as rylo-, or polyvinyl aczetal resins, Primers (low-viscosity solutions of adhesive dissolved in solvents) arc availabki for use in conjunction with modified cpoxy-filrti adhe~ives. Their primary functiorn is to prote.. prepared metal eurface-s rrom contaminatiop and oxidetion since epoxy films have adequate wctti~ig and adhesive characteristics without primers, Phensolic adhesive, 'iscd in the aircraft industry always are modificd &~ an clastainer or another resin Although they can be pruduced in liuid form, they now are used predominantly as films. Vinyl (polyvinyl formal or butyrill) phenolic adbesives wcre thc armt materials usei '9- aircraft metal boiiding. Rubber-phcnol.c evtisAi'ves include those modified With neoprene or nitrile rubber, the latter presently
)
heing the most widely used typt: of ela.ýtomcrphenolic structural adhesive. Lpoxy-phcrnolics arc used primarily because )f their outstanding temperature re!.istance. Being very i gid. they have good shear strength and creep resistance but poor peel and impact properties. 11ccaus,; of the poor %wruingarid flow characterisvcs of the clastomer-phenolic film%, a cojiing of liquid primter on the substrates usually is, advised. A% with all phenolic condensation rcactions, gases arecevolvcd during cure, necessitating relativel) high bonding pressures The actual pressure required to contain these volatiles is a function of thc temperiture: rise rate, 100 psi is a typical recommendation when the bondline is heated rapidly. All of the adhesive types discussed previously may he used for mectal-to-metal bonding. The selection will depend upon the relative importance of such factors as shear strength, peel strength, temperature resistance, chemical resistance, fatigue and creep properties, fabrication method, and cost. Generally, a modified epoxy or niitrile-phenolic film adhesive is chosen for primary structural applicatior.s, while at paste-type epoxy and simple contact tooling may suffice for secondarv structures with 1c, ;-critical requirenient%. The aequirenients fut sevefal cLa.,scs of structural adhesives are covered completely in 'IMM-A132 and M MM-A- 134. Although there is some overlapping. M MM-A- 132 is conceetned mainly with film adhesives while M MM-A- 134 generally has less stringent icquircments which are met by the liquid- and paste-type epoxies. Cured, reirnforced-plastic composites can be bonded to themselves or to metals with the same adhesive's and icchniqucs used foi bonding mnetals. In addition, adhesive prepregs can be used either for an !6
u.*p or ;.- L -odn"*-wer
ii,
.-
,-
--
e*-.-
a conventional r: i rforred-plastic layup and th'e substrate. These imstc~ials consist of a structural grade of reinforcement impregnated at B highi resin content with a resin formuiatior. having good ;dhcsion qualities. Reinforced plastics can be bonded to metals and other substrates by employing betweeti the substrate and layup a layer of conventional film adhesive that is cured simultaneously with the laminate. 1hi-; pirocredure is adv:'ntageous in that it precludes any mismatch of mz-ýtiqsg surfaces, a problem that always exists to somei extent with preformcd parts. While this technique has been found effective with a number of adhiesi',c and laminating resin combinations, such materials must be selected carefully for compatibility with both chtmical reactions and curing temperatures and pressures. Most of the epoxy adhesives also are suitabiz. for bonding facings to honeycomb core in applications 2-311
Ir
S
¶
whern good flow and wetting ability, and low curing pressure ate reluired. Some of the phenolic-bmued adhesives also may be wsed for sandwich construction, although most ar not reownmmnded for
through solvent evaporation rather than by chumicai cure, and therefore do not require temperatur or presure for curing. Because initial taCt often is adequitz to hold in position the parts being bonded,
this purpose due to poor lilleting action and the evo-
even clamping fixtures frequently are unnecemary. On
lution of votatiles during cure. When phenolic adbesivs arn used in sandwich bo-rdini, the core either is perforated or pesu is reased just prior to reaching the fral cure tempemture. MIL-A-25463 coains requirnents for adhlesive for bonding ssndwic•. It defines two casee Cila I for facingto-core bonding only; and Chus 2, for bonding facing to core and iQWerts, edge attachments, etc. Beause mos adhlesive suitable for sandwich comntnu2on also can be used for rmtal-to-metul bondin narly all samdwich adhesives are qualified to both MIL-A25463, Class 2. and to MMM-A-132. The adhesive puepdescribed previously also mnbe used in
the other hand. because these adhesives rmain thermoplastic, they lack the temperature and chemical resistance of tlh thermosetting tactural adhcsive. Where arnewhat stronger or more temperatureand chemical-rea•itmnt bonds are required, semistructural adhesives, such as the two-part epoxia and urethanes, may be uwd with room-temperature curing. Cements based up-au a solution of the polymer being bonded are L -ad frequently for bonding noncrystalline thermoplastics such as acylis, celluloWs, polycarbonates, polystyreves (including ABS),. and vinyls to thonselvea. The dissolved polymer gives
fabricating sandwich panel with reinforced-plaitic
body to the cement, while the solvent softens the ad-
facing. Out disadvantage of this procedure, how.. ever, is thkt a relatively porous laminate is obtaiad due to the lack of labminating pressure between cUl dub* Of SCVUwith Typical bond urergths obtainabie from several of the common types of adhesives ar given in Tab 217.
hecrnds. effcting a weld or bond when the solvent evapo-ates and the plastic rhardens. Transparnt acrylic patics also may be bonded
-5.1.2 Nemssmlra Adlhbes Thxse adhesives are used primarily to bond interior acasaor~s made cf a variety of materials, ineluding plastics, rubbers, metals, and fabrks. Because a joint failure would nor bc cawarorhic in thes ap-
plications, consideration of the highft- possble adhesve strengths is not paramount; and other fooon-, sich as cost and csanvetiemt. cMA be givcn tquaa matention.
"The adhesivte-. prefrred in thew applications
g'lrnelly are
baud "n
solutions or dispersions of
various elassomers and thcnnoplsuics. They set up TYPICAL PWPRtTiES
twnrftmwt
evllhoacnn.
t'nnt
A
clasioae-tesed adhestives are preferred; they offer oot adheson to mmmy matrals ad boetpr I -8
diffcnsat thermal cz iuon cosrAcieps. Elastoeneric
adhmves may be dn.ohlv in a suitab organic solvent or dispersed in water; tackifying reams, aetioxi-
darns, pca
~stiia and naforcia filles am uNWa
TABLE 2-17. o COMMONLY USED STRUCTURAL ADWESI•Y• T I-PEE LP
CHEMICAL TYPE
PIHYSICAL FORM
CURE TEMP."F
MODIFIE EPOXY 14YLOX-EPOXY
FILM FILM
250 350
5100 6100
EPPOXY-FHENOLIC
SUPPORTED FILM
NITRILE-PHENOLIC
SHEAR STRENGTH. psi -67' 75" 1u" 250"
5700 6500
290o 3400
1000 2200
350
3200
3500
3300
2900
NEOPRENE-PHENOLIC
F iL
SUPPORTED FILM
350
410
4200
2400
1800
EPOXY 'GEN PURPOSE)
2-PARr PASTE
15-200
EPOXY(HIGHTEMPMAODi EPOXY'HIGH PEELMOD,
I-PART PASTE 2-PART PASTE
25C 75-200
3500 I900
1500 3000 2000 2000 20D0 2500
1100
800 25W0 3000 400
OALUMINUM ADHERENDS TESTEC PER MUM-A-i3Z AT THE INDICATEDTEMPERATURES, 0
20 60 10
2 2 2
-
30 100
25 90
40
20
3 2 25
ALUUINUM CORE AND VACINGS TESTED PER MIL-A-25463 AT THE INDICATED TEMPERATURES.
S~2-32
SANDWICH
STRENGTH, III in. -67" 175" !80"
PFEL
iui.-Ib "31.WIDTH
-67' 75I
isO,
9, 35
60 170 33
35 45 31
NA
NA
NA
4
NA
NA
NA
2 2
NA NA
NA NA
NA NA
5S,
Il17
.
nOmi'~VI-.
mcthacrjlatc monomer ano acatalyst. whsctu have excellent strength and transparency. MIL-A4576 definft three typls of two-part ncrylic adhesives, type P contains solvent and is covered in MIL-P45425. Types 11 and lI arc withou solvent andmaybe u A for bonding plasics as covered in both MIL-P-5425 and MIL-P-184. For bonding of diimilar maerials, f•lxibk libmU Or fabrics, rubbers, or other such materials,
69mstrag,sjt; Skae~. 4-m W
ftj
36 ]z
(
,
*
" components of the formulation. MMM-A-1617 covers requirements for adhesives based upon natural rubber., neoprene, and nitrile rubber. Adhesives based upon natural or reclaimed rubber are suitable for bonding such items as rubber and fabrics to metals in applications where oil and fuel resistance is not a problem. Neoprene- and nitrile-based adhesives generally have greater peel strengths in the same applications, as well as good resistance to oils and fuels. The neoprene type usually is best for bonding neoprene and most other rubbers and rigid plastics, and has the best heat resistance. Nitrile rubber adhesives are preferred for bonding nitrite rubber, vinyls, and other flexible plastics.
)
Silicone rubbers should be bonded to themselves or to other substrates with silicone adhesives, such as those described in MIL-A-46106 or MIL-A-25457. No heat or pressure is required. Contact adhesives are a special type of elastomerbased adhesive having high immediate strength upon contact of the two coated adherends, but they do not permit any repositioning. They are covered by MMM-A-130. Other speciai-purpose adhesive specifications inude MMM-A-121. MMM-A-122. MMM-A-19, MIL-A-24179, and MIL-A-21366. 2-5.1.3 Pr•'silmg Operalihs Process and inspection requirements for structural adhesive bonding are contained in MIL-A-9067. Factors to be considered include type of surface preparation, control limits and methods of surface treatment. solutions. clean-room layup area requirements, prefitteng of parts, adhesive storage controls. handlin- of cleaned parts. application of primer and adhesive, tooling concepts. temperature and pressure controls, secondary bonding of subassemblies, rework, and destructive and nondestructive verification testing. Equal in importance to the selection of an optimum adhesive system is the selection of the best surface preparation for the adhesives and adhe,.ends being used. Some suggestions are given in MIL-A9067. Other recommended sources are ASTM No. D 2561 for metals. ASTM No. D 2093 for plastic surfaces, and Ref. 6. With some metals, such as aluminum. the surface treatment is practically universal: while with other metals, such as stain3css steel, it is advisable to evaluate different treatments with each combination of alloy, condition, and adhesive. Significant batch-to-batch variations in a given type of alloy may be noted. For most reinforced plastics, a wet-sanding treatment is recommended to obtain a water-break-free surface. When nonstructural ad-
AMCP 706-202
hesives are used in bonding, .i careful solvent wiping and/or sanding treatment wi:! suffice for many materials. 2-5.1.4 Desgn of Beoded Stnrctures Adhesive joints should be designed so that t~ey are stressed in the direction of maximum strength. Thus, the adhesive should be placed in shear while minimizing peel and cleavage stresses. Maximum bond area and uniform thickness should be provided for, and stress concentrations should be avoided where possible. Scarfing and bevelling are two methods that sometimes can be used to reduce the cleavage-stress concentrations at the edges of lap joints. Test methods for sandwich constructions are described in MIL-STD-401. while numerous other test methods for adhesives are contained in FTMS No. 175.
2-5
SEALING COMPOUNDS There is a degree ipoverlapping between sealants and adhesives: most sealants must adhere in order to be effective, while'an adhesive generally seals the joint that it bonds. In addition, many sealants are formulated from the same basic polymers that are used in adhesive compositions. Sealants are related particularly to the elastomeric adhesives, and many of the qualitative comparisons made in the previous paragraph apply to sealants as well as adhesives. In order to form trowelable pastes, sealants are formulated with hirher viscosity and lower tack than are the elastomeric adhesives. Lower-viscosity sealants also are available and are suitable for dipping, brushing, and even spraying. These materials. n•,wever, are classified more properly as coatings. Commercial sealants are manufactured from a variety of polymers, including polysulfide, urethane. silicone, neoprene, acrylic, butyl rubber, chlorosulfonated polyethylene, and polymercaptan. In addition to the base elastomer, a typical sealant formulation may include curing agents, accelerators. plasticizens, antioxidant-, solvent thinners, and inorganic fillers or reinforcing agents. Sealants may be one- or two-ccmponent types. All of the ,tter cure into tough. thermoset elastomers. The one-component sealants are subdivided into three categories: nonhardening putties that remain permanently s.ft: solvent-release types that become s.inihard through evaporation of a volatile ingredient. and types that cure by reaction with atmospheric moisture. Pruperties of the latter, after curing. are similar to those of the cured two-part sealants. (One further form of "sealant" is the cured elastoineric tape or extrusion. Because these nm.st be 2-33
AMIP 706-202
-
held in place mechanicall~y, they moreF properly might he called gaskets.) All effective sealants inust have a high ultimate elongation and a low modulus in order to acexpansion and contraction of the joint rcommodate being scaled. Most commercial sealants have these qualities. They vary widely, however, in their degree *of recovery, ranging from near 0% recovery (or 100% plastic flow) for a permanenitly soft putty to nearly 100% recovery for a cross-linked (cured) elartoamer. Thisproprtyis important because a lowy-recovery sealant. once compressed, must accoRmmodate subsequeni joint expansion entirely by its elongation, or it will fail. A compressed, high-recovery sealant will return, as the joint expands. to its original dimension before it begins to elongate in tension. Of the various chemical types of sealants. only polysullides. urethancs, and silicones are currently of I inmportaince in the aircrikft industry. Thcsc arc all high-recovery elastorners when cured. Polysulidecs are most commonly used in helicopters. where they act as both scalants and aerodivniasnc fairnng comvoundi. They have excellent adcharacteristics and resisianic to suicnis anad fucls, weathering and Aging, and icmrpcraiures up to 250*3. MIL-S-7124 and MIL-S-8802 describe the two-part elastomeric scaling compounds with increasiregly severe requirements for adhesion and resistance to temperature and fuels. MIL-S-87t4 compounds are formulated purposely with very low adhesion for such nses as fuel tanks access doors. A *grade for scaling electrical components is desciibed in NIlL-S85 16. Osme-part. noncuring. polysullfide puttics also are available. A mater-i.dl of this type is de* fiined by M!-Q-1!3 UR .1W ,it isitntfisr w ling ofitcal instrumenits, but is aseful for various purposes. Silicone sealants have autstanding environmental * resistance because they are unaffected, relatively, by * temmeatures ranging fromn cryogenic. to more than SW5*)F.~ and by moisture, ozone. and -iltraviolet radtat ion. However, because they arc the most expensive %ealant%. they are used4 only where these ex-cellent propcrties are requirod. Some: types also have very -htsion
*
gotod electrical characteristics, andi arm used to seal electrical systerfs. MIL-S-23586 covers siliconc scalants for electrical applications. and MIL-A-46106
describes a general-purpose, room-temperatugecuring adhesive-sealanit for both mechanical and electrical requirements. As ordinary silicones have relalively poor fuel and oil resistance fluora-silicone sealants should be ausd where these properties arc required. Both one- and two-component maierials mein common use. The former cure by absorption of atniosphem mc humidity, and. therefore, cure very slowly 2-34
in confined aircas or in thick sections. Primers usuaillv are recommennded to permit maximum adhesion to) m1cwKis Although of totally different chemical compost-Llion, polyurcthane sealants have many similtrities to the silicones. Both two-component and onecomponent moisture-curing types are common Primers (often silicone- based) are recommended, but. in this case, primarily fur retention of adhesion in humid or water-immersion situations. Thes escahnt exhibit complete recovery after extended outdoor exposure. They also are useful in cryogenic applicationr, where they are surpassed only by the silicones, and in electrical appl"cations. Polyurethancs also have excellent oil resistAnce, and grcater abrasion resistance than any other sealants- Osie problem is loss of adhesion upon exposure to ultraviolet light. An area where scaling compounds frequently are used is in edge-*ealing of honeycomb sandwich pancis. because joint expansion and contraction are not major considerations in this instance, relatively rigid sealants usually are employed. These are essentially thc same materials as the epoxy (and occasionally urcthane) pasic avhiicivci discuassed prev~iously-. cxcept that microballoon (hollow microsphcres of glass ( or plastic) fillers frequently are used to produce a 4 lightweight. closed-cell structure. Sandwich panels * also may be sealed with an edge w~apping of Fibergias prCsprCg. Viscous sealants mxy be applied with a variety of equipment, ranging from a putty knife to a cornpletely automatic mixer-dispenser system. Fluid sealV ants (coatings) may be brushed or sprayed. One.compawenst sealants suapplied in czartidges can bc apniod from miolan
ornl Ofir-oterated
*ums.. or these
seslants can be diwspesd directy from pails or drumis by air-powered Or hydraulic pumping equiptrient. Twvo-psar wsearlas can be. weighed and mixed by hand or by mcteing-mixing equipmient that dis.-el penses the compongnts according to prelet ratios. Frozen cartridges of pranixed sealant akso arc available4 commercially; thewe must bc stored at -40*F until just prior W~ use.
24 PAINTS AND FINISHES "IP
NTAD O IW RG 10 AITAN CO ItSOG II MIL-F-7179 prescribes in cletail the manner in which the external and internal surfaces of at khecopter arc to be finished. Uther helpful documcni~s arc TB 746-931-2. MIL-STD-171 (MCR), and AMCP 706-100. Helicopters require a Type I protection. ime., protection Against severe deteriorative conditions, F-or
241
)
most .;.jrfaces. this involve.; one coal of wash primer (MIL-C-85l4). one coat of primer (MIL-P-23377). and tv 3 top costs of attopcoat rot examiple, TT-E516 or M IL-C-8 1773. Preparation of the surface for painting will differ with the type of m, tal and with the surface (external or internal). For this handbook, exterior surface arc defined as all visible surfaces of an end-item that is housed within the helicopter and all visible surfaces of the helicopter. including all portions of the system that arc exposed to thc airstream. Interior surfaces are the nonvisible surfaces of an end-item that is housed within the fuselage of the craft. Prior to painting, aluminum surfaces usually are finished with Anodize MIL-A-8625 or Alodinc 1200 (N4IL-C-5541). and malnesium with Dow 17 or HAE (MIlL-M-45202). Nonstainless steels are phosphate-treated (MIL-P16232), stainless steels arc passivated (QJQ-P-35), and Fibe.'glas surfaces are sanded and cleaned with naphtha (TT-N-95). The first coat of paint applied is the wash pri mer. The term designates a specific mate-ial the( comd-
~hinr- the
-
nmrmrwflc nr shn inhihitiwi twacreth
.-sr.
craft arc the nitrocellulosec and acryiL-nitroluvllulosc lacqucrs. which contain a wide range of pigmcntatior. They art preferred becamusc of the case in iemoving thew. with solvents when it is necessary to change camouflage or color schemes or when repainting is required. They also are applied with a spray in volatile solvents (MIL-L-19537). TT-L-S 16 desceibes another suitable top coat. and one that as meets air-pollution regulations. This coating is a styrenated phthalic alkvd resin combined withi the necessary amounts of driers and volatile solvents. Thc mixture contad'ns 50* resin solids. inc~luding small peircenitagom of antioxidants, wetting agents. and stabilizers. A wide range of coloring pigments isV available, and these arc present in amounts of 24-45%. of the total solid conicrnt. There arc special paint formulations for camouflage. battery compartments, hilgh-temperature areas, walkways, and antiglare applications. Rain-crrosionresistant coatings (MIL-C.7439) are used on the Icading edges of the rotor and on radomes. There arc sixciail formulations for high visibility, and paints forC.f lettering and marking. Rubber, both natural and syn*hei,
...,,
surfwms suh as. ;!L-.
n
metal conditioaer. w~th those of the convcntio;:J-anticorrosive primer. The essential coDmponent)s D: wash primrrs are phosphoric acid. chromate merit, and polyvinyl butyral resin. Wash primers can be formulated that are effective equally over iron. steel, aluminum. treated! magnesium, copper, zinc, and a wide variety of other metals. The advantages of wash primers include ease of application and rapid drying, useful iange of tzmperature and humidity, application to a variety of metal, effectiveness in prc-
plastic windows, arc not painted. Particular attention must be directed to assemblies in which dissimilar metals arc joined. It generally is required that each of the mating surfaces shall be finished with the minimum number of coats required for interior surfaces. Where magnesium is one of the metals to be jcined to a dissimilar metal, tei metals shali be separated by MIL-T-23 142 tape or MIL-S8802 sealant. The tape shall extend not less than 0.25 in. beyond the joint edge in order to prevent mois-
venting underfilm corrosion, and good adhesion as a %peasmc suuuu SU"rI W&kla. 11 11U1 1eU4iUGLY
tur-s from bridging between the dissimilar metals. All naotes and couniersinas that attachin; pans pass through should be prinied, and all joining bolts, screw%, and inserts should be wet-primed when inserted. Preftrably. all stecl nuts, bolts, screws, washers, and pins should be cadmium-plaited. 46 PCA IW E
used wnsh primer is that defaned in NIIL-C-8514. a
smooth-finish, spray-type, pretreatment coating furr'ished in two part,~: resin component and acid comnP;ýwint. The materials must be mixcaJ prior to use. The piameir, which must cumform to MIL.P-23377. is used oam the wash pritm. It iscompatible with t usual =cYlic-aitrocelaloW Laqutu top camt. as well ass with the alkyd top coats (TT-E-516) sand urethane (MIL-C-81'.73). The two-cnsnpoximt, epoxy-polyaamide systern has high dhanaicu and solvent resintanc OWd Unusal wmhrbi.It is Wsjy-ajWIpled This specri~atic. also provides for an addkiicua dlams or materials wshitalc for usew~and ir-pollartion regulatiom. The aivailabiliy of daumas or costauiua Ai wpo~itgkg Smir rqpktimup is beomia# increasimigly iunponsat.. a=d this flactor shoul be kept in mind by the dein*=. Tb, top coatsw mofst osm pscifind for Azmy air-
)
42SPCA
3
FNSE
In addition to the finishing of surfaces with organic coatings as described in the previous paragraph, Ithere are a number of special finishes for metal which serve to provide the desired protection without further applicaiion of organic coastiags, or that arc used to provide a suitable base for the application of organic finishes, Many of these processes involve the devefopment of a durable crrosion-resistant. oxide layer on the surface of the metal. Although the development of this surface o.%de film may or may not involve the use of an clecuical curient. the chemical effect is similar and the proess is called anodizngs. 2-1)
-.
'
4 p
r
~ *'.'l .
I
there art many different finishes. For alan some or whic ame used to provide a bws for paint amd some of which provide protiection withont furthe painting. The processes all involve chromates noan eaidiAia iesgedimat, and rnll have proprietary compositioes. The performamot of these methods of tutanamn is governed by MIL..C-5541. The reagents say ba applie by sprying. dipping. or swabbing. generally. the metal is dippe in a sequence of baths ad inuus that cwrme dean, rine-rained oxide uniformy ditrbiated ovea the surface - with no coarse grains and so untrwated areas. Class I A treat. me"muim wihsawd exposure to sult spray for 166 hr and ane unpainte or follwad by wash prime and pimin tmtmeats. Class 3 coatings are simiLar to Chmu IA coatings, except that the deectrica reseisumie is low, Formagesim. hemarctwo primary anodizin~g manesium there ar te Do7tramn Foreamns (MIL.M-4S202) and the ote meho is the HAE treatment (MIL-lW43202). both of which involve elecroltican iziag f te mtalsuracein rde to buildolyui an fairlytikae of ahcmptlesrs=i ourder to a~
buid
~d
~
hic u
a lyerof aily
a..
''~
coplx
to provide hydro3gcn embnittlcment relief. Still another class of finishes frequently employed is the flame-sp.rayed type. In this technique, metals. silicon dioxide, titanium dioxide, alumina, or other ,m-A=rand -mnnr
aumium
is
'~
Ferrous metals that arc to be painted aregiea phosphate coating in accordance with MIL-P-16232. The~se coatings are of two types: Type M. which has a phos.phate base. and Type Z. which has a zinc phosphatc base. Type M coatings arc more resistant to alkalinc environments than arc Type Z coatings. When they arc applied properly, the reaction form?, a mixed-metal phosphate coating on the surface of the firrous metal that is dose-grained, Fine, and Wre of powder and course grains and that Ws¶the surface well. Thc treated surface is more res"5stat to corr-')3ion and provides a firm basc upon which to apply wash prime and prime coaings. In all of the foregoing treatments, the metals employed, especially the ferrous mectals, aro subject to the absorption of hydrogen from the solutions, and the hydrogen serves to embrittle: the metal. It is desirable to promote the diffusion of hydrogen from the metal by heating tt 210a_225*IF for 8 hr in order
fa,
-.. -..-.
n
ý... -
oxyacetylenec flame
~~ria or .Ai imln a - - --- -,;o.
nittnam nr -*. r
either as strands or powder
-
outdoor exposure without further coutin~s. The prIewhere they are vaporized and deposited on any subfetred method involves pretreating. primting, and art strate that wili condense and hold them. By this epoxy-polyamide finish. means, similar or dissimilar metals cam be applicd to metallic or nonmetallic surfaces. Ceramic materials Part made of co. rosion-raistrat sfteel arc passivated in oider to deveiop their corrosion-resistant can be applied in order to provide abrasive surfa=es qualities. This preceess serves to remove the "activewear-resistant surfaces, or flamec-resistant coatings. centers on the surface and to leave a thin, durable, M11L *674 covers the flame-spraying orf metals. transparent layer of oxide that prevents further corWo., a metal surfaces can be built up and subseauc oroxipaivhebyac ioesng the mea.parssinatn aqeos puii~s mtc.hOned uslapicain hade enai be thefam-stns rcomlshved or oimmierstaking the mertal ainatin iqeosquetc.:l mahned usflapicain orde toeepai thftpisons ?$AUSUu
t
nW
~
urnu
01Aumn
W
M.#inous20MUM
charomate. The temperaure of immersion varies from. 70' to I 55F, depeniding upon the alloy involved and the intended operatiag tempeature. Prior to treat ment, it isessential to wash parts carefully inan alkaline solution in order to remove all of the particles Owr iron that may have acewiwlataid ea the surface, as these woulod develop rua staie. durting the tieatnh.&it. The paswvation orocus maistatied in QQ-F.35. Ferrous surface that ame not to be painted usually are treated w~ith black oxide. The insulting depos*it is a hard. durable. oidin surfac. dma isattracive and is so. cwhat res"ISta to coroiomn aOW to wear. The process is applicable to both womstaolese ami simailes steels, it invoives imnaersiM the previously cleaned part in an alkaline. or alkafine-chromtate, oxidazing solution, followed by warmi and then cold rinses. a fInal chromaic aciod dip, and drying in
spraying oi toaimur
Iwater
worm air. The proý is defined in MIL-C-63924. 2-36
cauxvic contingi upunthe1 mflt5i
exhaust skirts or*je enginies for oxidation pl-3tection. 2-403 PLATINGi Another method of applyi-ig attractive, durable and abrasion- and corrosion-resistant coatings tiG metals and plasticsias meta plating. The platings of most interest in lieheopter dwmii are copper, nickel, chromium, and cadmium. Extupt for electrical cornpoms whecre electrical conductivity is impcrtant. copper plating is used only to provide a bae for thc wear-resistant nickel and chramijim plating.. Chromium and itickel plating. am used to provide hard and wear-resistani suirfaws for suiad object as SWa fasteners, strap holders, handles, knobs. seat arms, instrument parts. and other itemi where painting would no be satisfactory or economical. Cadmaium and zinc plating are employed almost exclusively to
provide galvanic protection Mgainse cxonvon. Cadmiuna is fthpreferred ecoming for ferous meal iAem
-
"
k2
such as nuts, bolt, screws, inserts, Lod pins used in assembly. particularly where dissimilar metals arc
izing plastics, and. becausc the danger of hydrogen
employed. A treatment of m( tal plating can be found in Rcf.
embrittlcmcnt is nqI1gi.', it also i- used for the plating of high-strength stmi: parts for high-stress ap-
7. In this pr
s. & ur methods used extensively: chemical reduction or elt-ctroless, vacuum vapor deposiluin. and molten metal dip. In electrolytic platin&, the item to be plated is cleaned so as to provide an oil- and dirt-free surfic, and then is connected as the cathode hn an ecoctrolytic coll. The anode is mtde of the plating material, and when an electrical current is passed through the
plications. It is the preferred method of cadmiumplating high-strcngth bolts and nuts and other fasteners, Fnd MIL-C-8837 detail. the requirements for this application. Galvanized steel products arc made by dipping the cleaned, preheated steel in molten zinc. Cadmiumplated parts also are made in this manner, and earlier tin coatings were applied by the dip process. How-
electrolyte the netal isdeposited upon the surface of
ever, the dip coating of steel with cadmium and tin
the itn being plated. In somi c4ses, the item first is coated with a thin layer of coppep, which adhers
has been superseded by the more economical and more prccisely controlled electrolytic procesm.
Selectrolytic.
readily to the base meta and forms a finn surace to which the platirg metal (either nickel or chromium)
U
•-
and tenacious coating. It is used frequently for metal-
MIL.T-l0727 covens elecroplatingt and hot dipping of tin.
can attach firmly. Ther are a great many proprieIn both the electrolytic and electroles processe, gtuy Oe•aolyuc sciutior, formulations and processes. hydrogen embrittlement is a ,agnificant hazard. DifConsiderabl, skill is required to ob.in a uniform, fusion of the hyIdrogen into the metI under the elecfine grain and brih coat, and much care must be trolytic forces is gSr-'ter than in the case of elctaoawcied to asure cleanlisms avoidac of poisom less deposition. The danger increases with the pn enUY of 41rms. biisenug ad cracking. an strengah and moduius of0t1 iatd materiai. 1-01u it avoidanc of ccarWraieed platings. Appliable is necessary to program a hjdrogen embrittle•mt t'CFedera Specificatis am (JQ-N-2l%, -Q.C32O, lief heating cycle in order to piomote the difftinn of Ja QQ.-P-416. hydroge: from tie basic metal. The optimum t,:mc LkElctr s or chemical-redutin plating depends and tempcrature will depend upon the nature of the upon the gena'mtioi of activated atomws of th1; metal coating material ant of the base metal, as well as to be deposited aiacemt to the wafAvaa•a eal kurupon the scheduling requirememts for the part.
fSX upo whicbh tht olatiMg is V.be m4xiAwJ. Thre are proprietary disa*
mCpo74a,• lsie5 u
lion presses for maya sia~n
rodtc-
used in pktling, and
prqimated or nonimpimqsate, and made from mnasy
the nicel re.1coig •se•v to make the surface more wear-resiwua mand kwa impact-misitive. MIL-C-
of the advanced plastic materials. The tapes asy or maly not have adsive on out or both side
2674 covem the p4Asat of eleciroless nickel.
(peswAre-sentsitive tapes). One application fo tapes is in the maiking of heli-
cop(t.
at so"
M IL-P-36477 may be usd isinlis of point for al rea.
Deca. confonming to the mquirments of
Plun4 wa*l, and are related st as to o,4tni a uii. form cmiosa- Tim son of the plating "4 may be
teusbal and internal mnArkings within the sim limaits spcimd. They may be prasaare4-asve. adhsive-
a hot war or a mohtea pool of t1w mN.a. Th metal is healed ekricaly W a taeperamtun a& which it veporimir, foui thLblsrfac. U ,tr dzh± & ,,t of an eketrosatak held aqp"ied betweenw iwtl swurce ard
backed, and scord and are aplied over pre•,AMly finished surfac. Antislip tale a&re umid on walknays, seps, mmd similar artea wULhiglrvn L" ate m• rapioyed iiiWn)C waica Wnia. Cautio. siwWbcid hausd
the iftsu biti plaed, the atoms. of the m"ga beCots se~amedsad sawatracted! to thme sekrfta. T'his
Pro-m jwodsii
wi ewapuicety bright, coheret.
,
I ..
K
., 244 TArt-I Tapes of varying ompoeitL'on. texture. thwknkos and width arc used in a varidey of ways in helicoptr. design. Fairic tapes may he woven or mmnwoven. ia-
vacuum dckibmr. 'hM i•,m to be plae arc racked of the
L
vamuum plating process and from the molten ametl
Such as nsickel-pltns of magnesium sdaafams whene
from th sorm
1)
drogen embrittlemnt eli•f. Embrittleasnct from the dip process is minimal.
cput•am dAat
14-V
Generally. the platinqpecificenion wil require hy-
also for Sot metl *&s plastc busc uateral,. Them• is less dwaola of hy•#d-,o evibnittlicnent with this prOOM. The' etolmas m9thod is unitablC for healird y;na or for us ift the field or shop wher Lhratoay or proce line facilities arm twn abailable. ft aho is ef~ctive for ==ne of the more diffcult jobs,
Va-uum de*,iton plating is conducsed in a
.
to ilatuiv tisef antisip tNpe edga arn rnot epusmii to ai"ow thatcant eta" the a to gaol. Thm - of 3-37
d
Myk:i
and p,1)yw~dic
tape.. in. lctruically driven
moxois b-?, aiwak possibt: the consxrv.ction of light, powerful motoii that cool rcadily. Tapes also are used for bindanM wiring haratýs*, sealia&accss pout3, and scaling cttities far foana-ir-pitaci filling. Teflon urfaces for tapes are used to pri&tw4t~ sins ebj. to skilin contact. low-loadl-bearing Finally, masking tkpcs arc. rvd in the finishing, repair, or repsaitina of surfaces.
2-7 LUBRICANTS, GREASES AND HVDRAUUIC FLUIDS 2-7,1
ENERALpropriate
-a 4
*
VJ
' ~~
Ncl;l~
Although this chartes is &'vo~ed to materials, the subject of lubricants and hydiaulica cannot be divorcoi frem; the total system. The designer ca'rnot Limpi) choose any oil or hydraulic fluid fm; a particular use. A lubrication systw. is designed for dclivcr-ing oil to the moclawnseas to be lnbr~ictct. Sclocrion of the oii to be. use is. as niucli mfunction or the filtcrnng. coofiag, and pumtping propenins of the.
Yima~eand brail;nlS
Ofw 1ar-tcwiMýu a a sj'itav.
raihcs thaethik.ivuh ricatiag cudai ctnhti"- friction iZ4, as *,I
'hi
lubricated. 2-7.3 GREASYES A detailed disco'ssinn of the many appro-,'co grrna is includrd in MIL-HDBK-275. For pur.poses of illustration, the applications for four types of. greases arm discussed. 1. ikclicopicr oscillating bearings. A sutitable
greae (1)1 us in bearings having oscillating motions of small amplitude - such as helicopter rotor head bearings - isdescribed in MIL-G-25337, and is apfor equipment that must operate at amWent temperatures of -6?*to l60F1. It also should be used for ball or roll"r bearings operating at higii Specn or hig~h teu'ycrstuns. 2. Ball and rolle-r bearings. These bcssingv. may be llubri~cntcd With Lga-as dcacibr.4 in MIL-G-2101 3. This grease is intcndcd toi use in the tczaperalure ravgr or -Iwo to 4509F. and is designcd pmtticulerly for those high-temperatUre bali and roller bewwri;4g applications wherc. soap thictenen mamy not
ay vwzatncvd 1<30,0W. Thii gttw; is not to bt: tist a4X 4p~calv t Lic 5-. care. Sm~vconucr~cmn tcia~on oSf hrdrmulic flivi&. Lben tWith gsa.tiŽ ti &i1Jwtt-nmt lus. SUlk 114 p'nrMAI rnirOf a~c rdwaa~sa8( thtM r~Io.VtL41GI lxaiagc, epu g-rs and gcar tfivwi; 3. Ccuw an d acvators. A gencrzl-pirpcLLz gx;evsv Of *16qutk~c filtih&s's buihzsiP. sId scals; m~us be coci~ , eato ~ otelbc~innc ~ ~ ~ iSic ' 41aee fo. v,ý&i;p.cer requiring a lubrian il.%hjijk lo;4-h grmses tiCW cjxccity ;isdescribed in MIL.C-9l322. Tiýis gan~sc is i1ese ar: dyrcamic raI.u than static uici kci uMs.4 in thn temaperature range of -65" ;o 3WIF, materials. The soimW range of lubricting rnateriais aud iii coinpstible with rubber. anid their QXAfltiO.S amt indicated itc labk 2-18& 4. Pzs.aiuna& systetr~s. Anothes froqicxntlv it-A Even wtvcrr th.. syý,.,imi is no dyr..mncx tand Manqumndi greavc is described by MIL-G 4343. Tb'* "shltapo4Ah- si lt &. w.o ne g1rcas is minted- for usc in pharumatic systwma a a t1K frctiad iw.0hVC4 mwtI hx s
-~
pcrforrn as rcquired far the specified interval when.so
iialatad cun.
of a C~ktcora &-zd a (Afr*,vrar$sdlba cation inizrtal for 2, girVui aPizsF&ainia ctbaritbtta as-
aurpAsc by lar ota~acwr thatr ncAi equ.pu'an will
j (
fnceo<. in secswts;. at, j::evfgretY.,tn psi. I nzk a&W for FIst wihb W'4L-P-55l6 rut'bia, but 142Cod noio be g&4. 'With alie( typ-M o08-bb7 withomt'. f-rtAli frc0snnpztflii-t
2ti D~&~% 0bL1J±~C~iGN SYTEMS 2-7.4 Wtif FILM AND PERMAiN1?T Th.; rcqueswwauu [cat dm~gn of lubrication Sys.-LBUAY tWwru for 1rttinx" ckwnagA in heheogner% arm ot wcaua e h4ThFý Ie~cl of rAnd and duwi con. dsscg&~da in lap; : e.twg for tAanstAnmimlaia'c frequicatly eac4:'Ovntered ini liclic'-~eJ opv~ratyuu;. ii u, advwnAnhasc'a to use nonw~cttimg WWbI.. ~ranw ~ rr~i~~b~a lbeiatio m~disuss4 i n ChajkA., 4. Af rjzasaittd*in Cla^t 4. tjjme wfttlcittM carill that do nujt auttvr thc duo ntotdizac. lIM{Adtd Md:ptn~ca ~piil lu~itsu~ wafo! apprQvi!a affr..1 m-a ul irel phsapic swatciialk (A lnj tcdi . al eiyt&j, hti~cl~xtK4, cpf u;lia mohv&&n. pruu-" to Ma..-,w ofi "rich l'kucA C%-vinn claca
1
i~fkint
haP-~ -- c nsakris. nod porous-eir
urcr.si¶
nT~nadnln
(uruxl~j-
~lbInudi
%L1.1 ot :ias-,6cnc. 1~ms aaea.shy arc enmptuyi4 Vinpvý ILo:4 tappf"Wkl~. icfx at it 10C1 nd Auiraa~ i en ~na&-w4 a"Uptitoij.
4
lubricant which arm intended
isapplied as a
isufide fcqu Molbdeum
phnolc o epxybonding luricnt n dryfil agent to provide a secre boadins to the mnetal baic. MIL-L-460t0 describes such a beat-cured. solid-rdall
vent piling and seizure of metals. Dry-trnl Iubrncants may be used on steeL, titanium. aluminum. aluminum alloys. and other metals. They are useful
lubricant, and MIL-L-46147 an air-cusred solid r~t,
NEU
where conventional lubricats ame difficult to apply
TABUL 2431. IOTrx LUSRECANTS AND HYDRAULIC FLUIDS CONTACT SURFACE
SPTEUS
GREASE
MIL-G-43113
PNIUMATCSYSTEMS
RUBBER -TO- ME TAL, DYNAMIC
MIL-G-?7617
PLUjGVALVES aor
FUEL OIL SYSTEMS
MIL-G-603
PLUG VALVES
GASKE TS, VALVES -GASOLINE AND OIL RESISTANT
MlL-.G-?1Ifl
tM" LOAD STEEL SURrACES
STEEL-TO- STEEL. SLIDING
MIL-r,4.12?
SALL ROLR MEDLE KARINGS.GEARS, ACIJATOR SCREWS BALL ATM ROLLER BEARINGS. -100' TO.850*r
MTLT-EA.WD
IIIL-G-25013
-7 ) LUSE OIL
to reduce wear and pre-
MEA-T-.L
EPRNE 9G
IETEPRNE
O
IHLA
TANGENTIAL ME VAL, ROLLING CONTACT
UIL-G-Z5537
I4ELCOPIER ROTOR HEAD BEARINGS. OSCILLATING METAL-TO- METAL. SLIDING- SMALL AMPLITUDE BEAR INGS
MIIL-G-8I3ZZ,
WHEEL BEARINGS. GEARS. ACTUATOR SCREWNS
METAL-TO-METAL, ROLLING AND SLIDING
MIL-L-357?
11IND1SHIELO WIPERS
METAL-TO-METAL. LOW LOAD. SLIDING
.IL-L-3918
INSTRWIE11T. JEWEL BEARINGS
STEEL AND JEWEL PIVOTS, TO-408F
MIL-L-6081
TURBINE ENGINES
METAL-TO-METAL, DYNAMIC
MIL-L-7806
TURBINE ENGINES
METAL- TO -METAL. DYNAMIC
AIRCRAFT INSTRUME.NTS AND ELECTRONIC EOUIPTMENT
MTLf-EA.FROSADNNERU
UMIL-L-6086
GEARBOXES. EXIREME PRESSURE
METAL-TO-METAL. HARDENED
MIL-L-2?36"
HELICOPTER TRANS?.NSIONS, GEARB0XES
METAL-TO-METAL. HARDENED. TO -40OF
MIL-L-22051
AIRCRAFT PISTON ENGINES
LIETAL-TO-METAL
IL
O5
MIL-L-?76941 TACHOMETER GENERATORS. GYROMOTORS, GIM~BALS CORROSION MIL-C-65?9 IAIRCRAFT ENGINES I'REVE NTIVE TIRE - TUBE. VOUNT IUG. DELO)UhTONG LUBRICANT MIL-L-836?
MAETAL-T O-METAL. LOW LOAD. HIGH SPEED MINERAL BASE OILS. 50 lit MAX OPRTO PRTO METAI.-RUBBER
SOLIP FILM
V3L-L-8937
CAIS.TRAflK(S.ROILLERS. SPHER~ICAL BEARINGS
LOW-LOAD SLIDING CONTACT, SAND AND DUST ENVIRONMENT
COMPOUND
MITA
THRE ADS.SS 9GL TS.PIPýNG . MOUNT INGS
ANTISEIZE. GRAPHITE . TO 400"F
IZAG14ETIC COTUPASS SERVO SYSTEL-S. CRANK CASES. GEARBOXES, FLUID TRANSUISSIGNF. ELECTRONIC INSTRUMENTS AND
ME TAL-TO-METAL. METAL -TO-JEWVEL
LANDING; GEAR SMOCK SIRuTS
FOR SYNTHETIC SEALING MATERIAL
LIQUID __ MIL-L-5020 DAMPING MIL-S-81087 FLUID MIL-H-5606
1 PRESERVA V~l 011. TFTING AN[I STORAGE ~IL-H - Iil) I11H'IG T! IPFRA IURE SYS TET1. 40' 1O &55', ~.TL-H-8IIRULTRA LOV TELTPERATURESYSTEL:. AUTOPILOrS.
IMIL-H-6083
HYLRUIC
WIL-H9328?AUTOPILOTS. SHOCK ABSORBERS.SRAKýS.SYSIELIS
MAETAL-TO-METAL. SILICONE. -10r" TO +5Wr
tdONOPERATING r: UIO
'.*INERAL O:L. PETROLEUM1 BASE
PC TROLEU'.* BASE FOR SYNTHETIC SEALING
rIRE RESISTA14T SYNTHE7IC HYDROCARBON BASE
2-39
or retain, or where other lubricants may be contawi-
REFERENCES
natcd edsily with dirt and dust. They gcncrally arc suitabk for sliding motion applications, such as in flap tracks, hinges, and cant surfaces, but may not be
i. Defense Metals Information Center (DMIC). High-Strength Steel - 18NI Mantrgahg. Battelle
used with oils and greases.
Memorial Institute, Columbus. OH. 2. Defense
Metals information Center (DMIC),
2-7.5 HYDRAULIC FLUIDS
Properties of New High-Temiperature Titaniunm
and data requirements for installation, The design, are covered by MIL-i-[-%W4 and system~s hydraulic
Memo No. 230, Alloys. Institute, Columbus, OH.
afe discussed in detail in Chapter 9. Two types of systems are defimnd Type i, which is designed fo, the -65? to 160"F rang; and Type II, which is used in to 2751F range. Two dasse of systans are the -653 defined: 1300 psi, which has a cutout pressure of 1500 psi at the main preusure con olfing device, and 3000 psi, which has a cutout pressure of 3000 psi. As in the cas of the lubricating oilds, alectioat of the hydraulic fluid is integral with design of the system. Pumps, motols, flight control actuators, heat exchangers. flexible connectors. packings, fittings, filters, accumulators, and electrical interconnects must be deshown to be fined carefully and their characteiti compatible %ith the fluid selected.
(
Battelle Memorial
3. Defense Metals Information Center (DMIC), Joining of Titanium. Report 240, Battelle Memorial Institute, Columbus, OH. 4. Structural Design Guide for Adwanced Composite Applications. Advanced Composites Div., AF Materials Laboratory. 5. L. H. Abraham and L. J. Lows, Shell Instability Problems as Related to Design, NASA Technical Note D-1510. 6. Charles V. Cagie. Adhesiw Bonding Techniques and Applications. McGraw-Hill Book Co.. NY, 196". 7. Heat Treating, Cleaning and Finikg. Vol. 2. American Society of Metals (ASM) Handbook.
I
zt.
"" F-
1
2-40
AMCP 706-202 CHAPTER 3
PROPULSION SUBSYSTEM DEvIGN 34 UIST OF SYMNOLS A, A.. -
area of inlet, inside diameter. in.' maximum inlet frontal area. ifl.2 ambient pressure pesi maximum VlOity sA01111 leading Wdg Sur-
face. 11)5
V, - inlet velocity. fps V free stream Velocity. fps M fuel flow. lb/hr AP -inlat pressureloCw. psi 3 -?P1l4.7. dinmensio~nlas
Ff
I.
-"
NM D CI 3.1 The propulsion system es defined during time pit liminary design of thme kviicopte after the enginc OT engines hWe been selecled, and their location in thme chsen Tb.chateris oncnal airfamehasbee ~ .- tie-.1the wishxew-1 syshthe 40con dmiratiof am sysem ofgrtos Th propulsio systm consists of the engine or engines. air ninduction sub.ystan. cahaust susstm fuel and lubrication smAbyztems Starting subs2yztc31. controls, tranaisassnn suabsystem. auxiliary power unit (if applicable) and infirared radiation supprassian subsystem. It also Ael inctud cooling an fir protecton subsystems. The air induction subelg also .u~frame. ie~ Ino "h parsahs ine peowa " coeideaaiss.proulson ontol equresemfs. uel and lubrication subyssem reqwrsawmts. compari mecat cooling. accessores. and AWU(auxiliar power unit) design reqiremeins ame discussd. The transmission subsystemo is discussed is Chaper 4. 3-2 ENGINE INSTALLATION 3-L GENEAL. iMany diffausiat engine installation arrangavients are pouible; frost drive. nor ve. side by side for mokit~iengie licoptesr. etc. Engine locations may be ssseiad so as to cotrild the overall CG of the beltcopter to okeai the momt effecivu powe train eon&Wunctmion well inmary. AU the mhysassesa dinually and topher a am inwingrl propulsio system. Emin hismnlasions usually faog within one of three thu complasy sunhaeqrdt inistalaion, the W.msmsm semi-exposed ing11alulsom and the exposed imialla-
umgdluah 31. The submerged engine installation places the engine completely within the airframe. This arrangement, an example or which is shown in Fig. 34I. requires careful considration to insure adequate accessibility for miaintenance, Removable firewall Sections often arc used, but this requires attention to the detail design of seals and securing dcvicca to insure that the fire protection of adjacent components and crew stations will not diminish with repeated removal of the firewalls. Although it may be difficult to service submerged
engines because of limited accessibility, their location often makes it possible ror maintevanc personnel to
work from &round level or from the cabin floor, eliminating the need for built-in service platforms. because the engine wsay be located deep within the airframe. engine air induction and exhaust ducts are greater in length and more complex than usually cyo tcrcon.titrtion. However, this inherent o auiio inefficiency may be offset by the extt real cleanliness of the aicat iib 312S~~ie Tesmepsdisalto sal mk h Thn be xpose t minsteallatox usallytn topo the tve buetwege. Theeniesmountgearo ad thetop of thear Teegieismutestrclaonhear This method. an example of which isshown in Fig. 3-Z. tequires only superficial structure to comnplete the engine enclosure. Teheh of the ern~vile ini this type of installation usually requires built-in service platforms. These pkatforms often are built into the cowling, so that the work platform automatically is available: when the cowl is opened. WIhen this isimpossible and separate platforms are proided, the engine cowling can be much lighter bmcuse it neod not be Structural The closeness of the rotor to the engine instalation requires that careful consideratioa be given to the cowlýlocking mechanisenit should be casily operabl by one man and capable of being inspecte from ground level for adequate security. The semivixposed instalation lends itself well to the use of ather front- or rear-dirive engines becAuse the engine can be located cither forward or aft of the main grorbox. The rer-drive en~ine stay result in very sbon cogine air inkdcton subsystens that can be anti-iced officientdy by wart hearing SystemsFront-drive engume may resul in snore complex induntioo subsystems. sad. because of ther comn-
' 4
4
4
d
VA
Figure 3-1. Subme~'d Eaglet Iuatallatioe tExample)
3-2-
pltcated shapc,.. may require more coniplexj anti-icing subsysiceim. Twin-enigine hclat.oplcrs often' employ the warnexpicdcofi~iraio. Tisenabknk a single housing unit to enclosec all engines, &Md p-rovades good - ccsai-;iihry to all cmagines provided ahey arc spaced su~ficieritly far apart. Additk~nal accc4.sibilitv can be ohtamned by making irierengirre 5ýrcwalls removable. If
side the airframec and 512 exposed on all sides. This arrangement, an example of whkt is shown in Fig. 33. commonly is used with a strc~aanlundc niacel that provides environmental protection and teduoes ftirodynamic drag. The externally mounted nvjsc atrangement provides the hest acwuaibihity. proviided adequate serv~cc platforms or other convenient work areas are available.
this is done, the designer must pay particular allen-
The nacelles sometimes are attached directly to
I-on to the seals on tin' reinosablc sections. Labyuinth seals are superior to otlier wealing melhtxs
vided by removable paucls. The engine mounts may
Intechagcailiy cjgin intal~iin btwee
may eteral urf-w my cnsit o higedpanels.
conskdcrable savings in initial and maintcniancic cog.s. along ivith increased aircraft availability.
loadsharing structural members of the nacelle. The pancis also may be designedJ w serve assa work platform. Thi~. arrangemitnt requires a minimum amount
jn11
aiiuse btallation. engines are located out-
fuselag frames. Acces. to the engines wenally ispro-
'ihity o! the eng;ine.
-3-
ANCIP 706-202 3-2.XIA
D
Checlist
3-2.
The following items are applicable to each or the preceding types of engine installations and, as such, are basic design objectives: I. A properly designed engine enclosure shall be: a. Aerodymanically clean b. Sized and proportioned to the engine and its related subsystems c. Fastened to the airframe, not the engine, This eliminates problems with metal fatigue associated with engine vibrations. d. Arranged in such a fashion that the major portion may be opened quickly for inspection and minor repairs, or removed entirely for major maintenance tasks and cowling repairs e. Adequately ventilated to prevent accumulations of gases. and designed so that accumulations of dirt, waste, or fuel may be observed without removal of cowl sections f. Properly drained so that no fuel is trapped in any ground or flight attitudes. Any fuel likely to leak into the engine compartment must be drained clear of the helicopter through an appropriate drain system. 2. Appropriate firewalls must be provided to contain fires within the engine cowling or nacelle, 3. Daily maintenance aids should be incorporated if the configuration will permit them. These include work platforms, inspection accesspanels doors.in and supPorts to hold cowling or nacelle an open position to hase maintenance operations and protect equipment against accidental dama ne d equMaxipme a insterchacc enabliy d pag. s2.
ENGINE MOUNTING The engine mounts shall be designed to withstand the loads resulting from the engine torque, thrust. and gyroscopic couple in combination with all applicable ground, flight, and inertia loads. In addition, engine mounts shall withstand transient torque and crash load conditions. The engine mounts and supporting structure shall withstand the inertia torque resulting from sudden stoppage of the turbine rotors combined with the flight loads for 3.0 g flight. Torque decay time histories shall be determined by analysis of the engine characteristics, but in no ca.e shall stoppage be considered as occurring ?n more than 3.0 sec. Engine mounting requirements are specified in MIL-E-8593 and shall be followed. Turboshaft-engine-powered helicopters may require critical alignment of high-speed shafts. Ii is good practice to design a high degree of accuracy into the mount. and supporting structure, and thus eliminate the need for adjustment on installation. rhis require,- more intricate tooling during manufacturing, but insures positive shaft alignment. For multiengine configurations, interchangeability is desirable and can be achieved by designing the engine mounts so that common detail parts can be assembled to result in opposite assemblies. The various of engine mountings may be described brieflytypes as follows: I. A three-point-suspension type that incorporates a gimbal or ball joint A mounting that cantilevers the engine from the
4. Maximum interchangeability of parts shall be incorporated into the design. 5. It should be easy for an observer at ground level to determine that the cowling is secured properly. 6. The enclosure should have easily-accessible provisions for fire extinguishing, by ground personnel, during engine starts. S7.All portions of the cowling that might be sub-
gearbox. Few engines can be cantilever-mounted; consequently, this method will not be discussed. The front mount of the three-point suspension may be either a single- or two-point configuration. The gimbal arrangement likewise may be a single- or twopoint support. The ball joint arrangement, on the other hand, must use the two-point support to obtain torsional restraint for the engine. When the three.
jected to exhaust gas impingement and to exhaust flames in the event of an exhaust subsystem failure, shall be corrosion-resistant steel, titanium, or other equivalent temperature-resistant alloy material. The
point support is used with either the gimbal or ball joint, it must provide engine freedom for thermal expansion in all directions. This is accomplished by providing lateral, axial, and vertical restraint at the wo laterally disposed points, vertical restraint at the single point, and torsional restraint through the two laterally disposed points. Other configurations using the three-point arrangement may place the gimbal or ball joint support alongside the drive shaft instead of concentrically as is commonly the case, Because the engine support and drive shaft no longer are concentric, a simple "trailer hitch" arrangement may be used advantageously for cost savings and to provide for easier drive
material selected shall also be determined by the heat transfer analysis considering the engine heat r tion. 8. Cowlings shall not interfere with any parts of the engine, its operation, its accessories, or its installalion. 9. Cowlings .hall be designed to provide udequate cooling of the engines and engine accessories during flight and ground operations. 3-4
shaft and coupling inspection and overall maintenance. With this configuration, it is imperative that at least one of the drive shaft couplings be capable of providing adequate axial displacement. The positive gimbal or ball joint may be replaced by an elastomeric element that supplies vibration isolation in addition to the flexibility of mechanical joints. Isolation mounting systems are discussed in the paragraph that follows, ENGINE VIBRATION ISOLATION Chapter 8. AMCP 706-203. specifies that an engine vibration survey shall be conducted to determine the
3-2.3
dilona lanstell nd prepared, and ground and shalltstbee pepaedandgrond dition. a test plan flight tests conducted to verify that the engine vibralion environment is satisfactory. design will require a flow of Successful the engine manufacturer. the airframe among helicopter data danufacta a r. ng the engine ma ingacture The . afre s
manufacturer, and the procuring activity. This flow S
)
data are defined in Chapter 8. the required and andthe required dtype
As pointed out in AFSC DH 2-3. a mounting subsystem shall be designed so that the natural frequencies of the engine(s), when installed in the helicopter, do not exceed a certain limiting frequency in those modes of motion that may be energized by the vibratory-forcing functions generated during the operation of the helicopter. The natural frequencies shall not exceed 70% of the lowest frequency of the forcing function, 3-24 IFIREWALLS To provide for isolation of fires, zones that contain both combustible material and a source of ignition must be defined and shall be separated from the rest of the aircraft by firewalls. The firewall must withstand a 2000°F flame for 15 min. Sources of ignition may be hot engine surfaces or electrical connections. High pressure ratios and increased cycle temperatures have made virtually the entire engine surface an ignition source. Consequently, the practice of defining the entire engine compartment as a fire zone has evolved. Stainless steel, at least 0.015 in. thick, is the most commonly used firewall material. However, in applications employing a structural firewall, improvements in weight and cost-effectiveness may be realized by the use of titanium or other suitable material. In such applications the structural requirements usually are predominant, and the material thickness required is easily capable of providing the necessary fire protection. Firewalls provide the most effective protection when they are kept free of sharp protuberances such
as angles. clip%. and brackets. This allows the fire-estinguishing system to operate more efficiently. Engine installations incorporating nacelles sualy require that only the interface to the airframe be ffreproof. This area, therefore. should be kqx to a minimum to achieve minimum firewall weights. Side-by-side ertinc installations require a common center firewall, which can be made removable to enhance engine accessibility. When this is done. care must be taken to insure a tight-fitting. rugged seal. All-metal seals appear most attractive for this application. Pliable seals, either butted or lapped. eventually deteriorate, thereby reducing the firewall the removable provides Thb. seal integrity. with a certain amount of inherent support,section facilirtating removal or installation. Side-by-side engine installations are not desirable due to vulnerability and sirations. sualion survivability considerations. On each face of firewalls. and immediately adjacn hrtuesol emd fmtraso cent thereto. use should be made of materials or a that will not ignite as a result of heat transfer from flame on the opposite side of the firewall. Combustible fluid-carrying lines that traverse a firewall shail be equipped with shutoff valves. 3-2.4.1 Fire Detectors Three basic types of detection systems are used: infrared, continuous wire, and spot (thermal sensors). The infrared or surveillance fire-detection system provides extensive fire zone coverage. AContinuous wire fire-detection systems are of two types: those in which the resistance across a eutectic salt filling an annular space between two conductors is monitored continuously, and those in which increasing pressure of a gas trapped within a sealed line pneumatically actuates a switch. Each of these types is routed throughout the fire zone in the areas where temperature changes caused by fire are likely to occur. The continuous-wire elemert is subject to vibration and maintenance damage, which can result in false fire alarms. However, continuous-wire systems are not vulnerable to false alarms from sunlight. The spot type of fire detector, or thermal sensor, actuates a switch to trigger the master fire-warning circuit. This type inherently is more rugged than continuous-wire detectors, but has very limited coverage. As a result, spot detectors in reasonable numbers can be used only in fire zones oý limited volume, such as combustion heater compartments. MIL-D-27729 covers volume surveillance types of flame and smoke detection systems. MIL-F-7872 covers continuous-type fire and overheat warning systems. MIL-F-23447 covers radiation-sensing (surveillance type) fire warninp systems. 3-5
3-2.4.2 Fire Extinl g ii Almost all recent hbicopter designs use high-rate. discharge fire-extinguishing systems. Most systems use vaporous extinguishing agents propelled by a dry charge of high-presstzre nitrogen. More recently. some extinguishers have used pyrotechnics as the propellant agent. Inert agents. such as bromotrifluoromethane or dibromodifluoromethane. often are used because of their good extinguishing properties and low toxicity. Furthermore, the low boiling point of the agents facilitates vaporization and distribution within the fire zones. An effective fire-extinguishing system is one that will. by test, demonstrate 15% by volume agent concentration within the fire zone for a duration of at least 0.5 sec. i'he system must meet the requirements specified in MIL-HDBK-221. 3-2.5 ENGINE AIR INDUCTION SUBSYSTEM
requirements are met. the duct pressure gradient i• w made favorable for the flow by decreasing the crosssectional area of the duct along its length and by contouring the walls of the duct to polynomial equations.
Two basic tools are usefui in the aerodynamic design of engine air induction subsystems. These are: I. Analog field plotter, which uses an electrically conductive paper and is based on the fact that LaPlace's partial differential equation is identical for an electrical field and an inviscid fluid. This technique yields local streamlines, velocity potential lines, and surface velocities, and is well suited to two-dimensional problems. 2. Potential flow digital computer program, which uses the technique of superposition of sources and sinks to yield the same results as the analog field plotter, but with greater accuracy. "Basiccriteria for the aerodynamic design of the air
where V, = inlet velocity. fps V. = free stream velocity. fps Beyond this range, the possibility of lcading-edge velocity peaks, and hence flow breakdown at the nose. increases greatly. Ref. 3 shows that a certain minimum frontal area is needed to keep the external maximum velocity within limits. This criterion is satisfied when
induction subsystem duct, which must satisfy the requirement of the engine model specification, are:
3-2.5.1 Air .dmetliou SWuiswu Deuign The external lip profile is established by fitting an external cowl contour (usually a NACA Series I or an elliptical shape) from the lip tangent point to the inlet envelope boundaries. The inner lip shape usually has an elliptical contour, and the design parameters are given in Ref. I. Class A Kuchemann-Weber circular intakes described in Ref. 2 yield design parameters similar to those given in Ref. I. Ref. 2 also suggests a desired design range of inlet velocity to free stream velocity ratio. i.e.. (3-1)
0.4 !_cV, / V,,_< 0.65
+4( I -
=- Ž I + A, (V,
V/ V.)
(3-2) 1
/ Vo)
-
1
where Am = maximum inlet frontal area, in3
A, = area of inlet, inside diameter, in.'
V,,r
= maximum velocity along leading edge sur-
i. The air induction subsystem shall prevent any erratic or adverse airflow distribution at all operating conditions and attitudes. 2. The air induction subsystem shall have minimal aerodynamic losses. A 0.5-1.0% pressure loss should be attainable in most air induction system designs.
face, fps In addition, pressure measurements should indicate that variations in inlet total pressure, evaluated in terms of a distortion index, as defined in the engine model specification, are within the required specified values.
When a particle separator is installed (see par. 32.5.2), the pressure loss will be higher but should not
3-2.5.2 Inlet Protection
exceed 2.0-2.5%. Each 1%of pressure loss results in
The engine air induction subsystem should be
1.5-2.0% power loss.
designed, to the maximum practicable degree, so that
3. The air iqduction system shall meet the minimum acceptable engine inlet distortion limits as pre-
foreign objects from external sources will not enter induction subsystems. The level of protection re-
scribed by the engine specification. The local total pressure should not differ from the average by more than 5.0%. Items 1,2, and 3 are interrelated and pertain mainly to pressure gradients determined by the duct area distribution, duct wall radii of curvature, and changes of duct wall curvature. To insure that these
quired for the engine air induction subsystem is defined during preliminary design. Various engine air particle separators (EAPS) are dccribed in Chapter 8, AMCP 706-201. An engine air inlet sand and dust protection device, if installed, shall meet the criteria specified in Chapter 8, AMCP 706.201.
3-6
An intation anti-iciag suabsystem dM11 be din-
if an ice dekector is installed in th induction confornm to M I LD-S 18 1, Any failure of the anti-icing countrol "hIl result in the anti-icing wbaylsyttu remaining in or reverting to the anti-icing ON mode. Eugiae air induction system~ can be anti-iced either elnctrically or 6y time use of engine bleed air. The former telncricaily) can use a nonmetallic duct in which thermoelectric beating elements airc embedded. The latter type (bleed air) has used a metallic duct that is formead in%* a double-skin heat eatchanger adjacent to thec area requiting therm~al protection, These bleed air heat exchangers have been made with and without fins. The decision between electrical and hot air systems is made for echb helicopter on the basis of the renalired heiconter misi~on and results of trade-off studies in which the airframe/transnuission/cnginc match is considered to determine the system which results in the lowest aircraft pealy
system. the detector slu
-
)
3-2.53.1 Elect"~a AM~e-idg Typically, electrically anti-iced helicopter engine air induction subsystems require a variation in local power density from 4.0 to 16.0 W/in.2 to account for local vakiations in surface velocity and moisture ianpingement rate. Relatively large amounts of clcctrical power are nececssary, which results in a substanI.-3
Lg
ntSp~
.-
r
weight. Electrical systems arc relatively easy to design and test. The surface temperature of the air induction duct normally is held to 40 0 17(4.40C) for the atmospheric dcsign condition the anti-icing system is required to meet. Calculation procedures arc contained in Ref. 4.
) 4
31-2.5-32 Bleed Air AtiI-Icift Hot-air-type anti-icimig subsystems use compressor bleed air, which must bk adequate in quantity and temperature to metl all rcquircmcntr. throughout the power and environmnental spectra. An advantage of t;ese systems is that the related powci penalty is ap~plicable only on a cold day. when anti-icing i% required. In most cases. the helicopter will not be ewer-limited on a coW day and. thmerefore. will surfer only a fuiel consumption penalty from the use of compressor bleed sit,
problems. howcvcr. Fins, if used. mums be braand
turing and aisainbly capabilitins also becomne important design considerations. In a bleed air anti-icing subsystem. engine bleed air is ducted from the comprissor bleed port to a solenoid shutoff valve and then into the intake manifol ait the leading e*g of the inductionl iibsyitrni inlet. The flow then impinges ovu the inlet leading edge. providinS the greatest heat transfer at the emtemal flow stanation point where the thermal Wod is highest. The now then passes through heat exchangeri alont tihe inner and outer lips of the air intake. The gap height (minimum gap - 0.000 in.) along the induction systemu flow passages is tapered so that the exsternal skin temperature is maiimained cloac to 40*F. The air then is dischar~ged overboard through discharge slots located at the rear of the outer lip. A thermal switch should be cusd to monitor duct skin tgm~p=c:*rc. This swtitch astu*tecs a warnin light if the skit. temperature drops, below 401F while the system i. in operation. Actuation of tbe light indicates ci~her a subsystemi failure or icing conditions more severe than the subsystem capacity. .. 3 idIlgDiusrte eesrl W-ce 3"3 The capability of the anti-icing subsystern must be demonstrated by test. The test requirements are described in Chapter 9. AMCP 706-203. 3-2.6 EXIIAUSY SUBSYSTEM In
Aý_g s..c
.,f h
ene n
ex-a
..
sy-.-.msalme J
the following objectives: 1. Minimize pressure loss to reduce engine power loss. Losses usually caii be held to a pressure loss of IlkY or less. A I%pressure loss grrnerally results in an approximate i% power loss. 2. Prevent loss of tail rotor efiacicncy due to hot exhaust gas flowing through the tail rotor 3. Prevent loss of power duc to heating of the inlet air and/or reingestion 4. Prevent overheating of the adjacent strv-_.turc due to impingemenit by exhaust gases 5. Provide maximum possible thrust recovery. Exhaust system assemblies usually are welded or furnace brazed. The brazed assemnbly offers high resistanccetic tal fatigue becausc the strcngth of the material is not uffected appreciably by the brazing operation. BRazing has both advantages and disadvantages. It requires hightir initial tooling espenditure%. hia%a potentially iower unit price, and usually is
haust cjectoý %as air pumping devices. Although many ejector configurations are possible, two cornmonly arc used for this pupoe The frsnt confguration p'rovides ant abinulus for momentum exchangr at the downstream end of the exhaust duct. The nacelle that forms the outer surface of Owe annulus extends beyond the exhaust duct to provide an adequate mixing length for this ejector.
pouuibility of capoure tolIL seekiaig4 usivpuxs an IftY radiation suppression subsystm awys~ be ~e The extent and type of suppeuason rmwývd will havc bee &flined by the procurngl acfivwtý. and inJoaded in tile prelimnwary design. As cdssi ad in Chi~r 8. AMCP M&.201. thUe IL suppeesioa subssystgo vatay be a part of the helicopiter or be a meprate ka. In *jw cm, thse suppresson of 1K ruia*Aak requira reducing the tm"Perature of thx tank sou te. The paragraphs that tIllow diacxts cal) passive countermeasures to ER weapons. Active cow~tcrmeasures also mny be rwquired but. as disuctexd in Chapter 8. ANICF 706-203, Ii.c quAlification of smch systems by US Army Akr~ation Systems; Command (USAAVSCONI) is Winiedi to tbS iaaa~cc bItwenn the leicopwe and the sabsyst~sn. The principal heat scurms are the vbg~ine hot part. exhaust duct, and exhatost plume. However. other heat sources also may produo" sinirscant =aunts of energy in the ER frequency band. 'fl. raiaition from other sources (ecg., heat exchaffW. r outlets and solar ref'lection from windshickti. also may have to be reduced to bring tht total IR signature of the tieli-
PICCtvcn dnasm 4#Cti 4are MUse Oe WILM LFu
... Ct-1 .........
11M esiffw*u t0 rePAir. BY cMUtas, W"ld require Iba goohng and Mge rqmpAie mome easily. buit are sumw ProIw to M~etal fatiue becaus of fthaetallurgccAnge in ft WeOd area. To reduce pmressurloss and thus obtain maximusm efficency. the zexuse systems skoul &oA make abnWp cross-sectitalchanges. Duct beaids should be gradua anid an adequate diffusion aesgle must be "ma rinted. In mQicrgine helioptes, savings iRniitWa eoat and greater availability or spare parts makeh imaserckaable exhaust duct- "esrablc. 3-2*61 Ihme Ejecters Engine compartment/engine component zooling is discusse in par. 3-6. However, engift installations requiring
positive compartmetnt Cooling Often use CX-
114X1
through exhaust gas impingement. Flighitemperulure liners often arc used in this area. The exhaust duct isof conventional design, mounted directly onto the engine, The second configuration locates the momnentumnexchan~le annulus at the upstreamn end of the exhaust duct. ihe engine serving as the inner ring of the annulus and the duct as the outer ring. This method allows
I.rates
the annulus rings, resulting in higher operating efficicacies. Two basic dzsigns are possible. One sepathe exhaust duct from the engine, thereby reducing metal fatigue of the exhaust duct, which is supported by the nacelle or airframe structure. The other uses standoffs to mount the exhaust duct on the engine. The latter must be used when the engine installation requires a exhaust-duct-im posed load -ýo change engine natural vibration frequencies. Generally, the selection of a method for pyoviding: the necessary cooling airflow will have been mad during preliminary design. As discussed in AMCP
706-201, par. 8-7,41, this selection requires cxaminL. lion of th%; wtight penalty iand pi,.er lessa ssociatted with the alternative means. Design cons'derations pcrtin.:nt to the integration ofm exhcaust Cje..tCT iI'!6 the cool~ng system are reviewed in paf. ' 6. lnclude6 are refrncrrcs to design proct.4urrs and rcquircmcr.ti for desijn. documentation. and demonstration. 3-2.6.2 Inidrarod 11tiR) I&AdIaOm It the missions assigned to a hzlicoptcr tnclu':-. he
3-9
uuErit
3-2.6.2.1 IR Sappresela Reqsktaflutti Thc I R supprmsion requirements for a nec* Army helicopter will be provided by the Syseiem Specificrtion. Typically, the requ'rernent will be *tAited it) the example that follows.
~IR SUF'FrkSImON REQUIREMENT The maximum total ;R railiazion siulature of the helicoptei shall be suppressed to lewels nct to exceed ý (2 desired~) W/Isr in the 3-5 mici-oun bandwidth. The total I R radiztiob! signature is comprised of d*rcct (i.e, ,4iihle hot metal parts). indirect or re.1cted, and cxha!Ws Vpuma rae~ation. Radiption frumn thr engine and tailpipe and from all secon~dary soutees, such as hcat exchanitei ottelts anrd fuselage areas washed by exhasist, aic in_;uded. Thc suppressed IR radiatknr s'gnature itqi~rement.; shall he based upon the Army Ho~t Day Atmosphcre (i.e., l25*F at tiea lriei) and shOl/apply will the engine o~eaiing at intecrn edizre pow~er and with the heiicortcf at the grsi wvegh: thmt rcsolts in the hightest %exhaustgas temptyature. The si:(natz.- sholl be rwiluated at lower hcrr:_r~'er. upper hemispht~rc, and coplansa' viewing argies, und the required level -.f supp:ession shall apply to the viewing angic thak results in th inximurnsigiiature.,
I
Rtequirasc for &iisfit ipociki sfnats sepleasuc so shec umainms accqpuabhe
I& uuppremaon "Wuufss so will be prov"ide ma ceoimiemi systam specification. The sipstae will be described in
pAsc ired(or tharW risedwo sath ea 1osevera Wevwhaoba wiýis she lIt uP@um
A ceutpresive WUstakai of military IR snckedgy itg&a by AMOI T06-t27 and AMCPWMb 922 (Icfrsud M~AwV Span fiPan, Ow and Tva. Rboo.Adhusme&l4WWO pelinet oue fm so Ill. wpphwowi~ls uaqprmmaue.m art gave by this two-pant handamdauiam will be p,,ovi~sd by USAAVSCOM tapest ftqms. Ilowever Compliance With apuo lequirwinmnt. will be dmnoustraeat by a IR ugjnalsare mLszvy (am per. S-t. AMCP 706-2W).
*
flU.Z Luum &%-qemn As nose prvissly. fth principal sos'm of It, ra4iatizu from a helicopter arc she visibl hot parts of tlit regime and Uslaggiac exhaus. Therefore she mons LmpOWPrtm pesOf shell suPPiMkO Mubsysnnoraliv as the eahamssmumujwr. Dwemndias ova It4 Mereica sumi so tWs hsam . It6 &upprcssor may be an isuwra pern of she helicopter or is Sue bed4ri4 as askit to bec installed only when she belicor. 4zu insacombat role in which epipagemet by iR setkir., weapon(s) is probable. Further. the exhutg IR suppressor nary bc dtvcluqa by the enginc conrtctor ard provided as an engine accessory, or the hcicoptew manufacturer may be responsible fur shg devvlopmn~rt of the suppressor and of the instatla-
MiTson
tn.rotor
Desir of an cxhau'st IA supprevior is a complex nrocrrn.The .tg
onntiuptrnajslihnhra re disrnwdM in
Chapter 8, A MCP' 706-20 1, mnd the trade-dft amiong shc possible vvrwer lasme and weight increases aluo are discussed. The methods of cooling a suppressor arc descrbegd and she possible use of low emishii ty miatktgs a~o isdiscussed. Further explanation of tht kelt transfer prooesses imacluding pertinent eq~uations and values for the applicable material props-sties and ohrconstantls, is givutz in PSI. 4. Typkally. she first requirement of an exhaust supprcscu-or isto shield she hot engine parts from external
V
viewy. The shield in turn isit*@lf heated by impinging
4Wtwl -
V
exhaust gab. The effictiveness of the shield as an IKR radiatos then must be reduced by cooling or by con, of she cmissivily of the surface of she shield, or both. H-owcver, she temperature of she exhaust plume can be h-duced only by coolirng and this really can be accomplished only by dilution; by miting cookst air
ami mqsimnmua she amassm of cedin air etAp ha quie larip and she po i rrequined to prowideoh viomary uirfic' "le SOnu -w a. hi. S ductibs.satud f atwv mijm/nune strigmas It mnppawuaem reqtiresmm with maiwmam impact "Mue proulion ystem 65lo-n pow wequred for ecitle mimem pwftormnm. Wbherat eah"msslo aqausoresusama&auian-A jpat part of she engine iaasalhstioa or a removablk kit. enine oufshe huglbeopser Futhe.em s isslly VAI m mot alew ademiely sUn vibmaom eniommiw4a#of se of an *%hase It sappressor onabelicoipsor "~t arv remilt im a reduction of e-a'n or engine osapari -Ln cooli% kuhn mama. sequined L"&l sands any nomtIna operating mmnditii.. Ahso. thininstalhsbion dd! no rodues appuecaify she tweet availabW from she menie, or wceuem apre.&mbl eithe Ws pwowe requird frogs dhe ngin or the W< baedl or. power required. Typically, a pow" loss not inesamb of 3%of she power fequird so howe as, the upogfis hot-day Coo~dision (C49.g.. kC t 9?F) is a0001116ble, athkous the lam aowable in a given case wil deuporn tin. mwwais" knL M ap5$ V"b. W flj 3cas, die iaatallation of any device to suppree the IR sin-r ala rvn h coinpt~iaace of she balicopter with she performance requimwaents of the system specification.
1
-=
3-3 P ?ULSON CONTROLS
ContrcJ of abhelucopter p-ropulsion system rcqwrea considerution of sht. churactristics of she hidkoptcr as wvcl! as thle Ut.Th oaxatrl system is defined during prlinainary design, and the ....
Z...a.f.
iE aAeaiI
.-
in Chopwr R
A 7621Thdealeinegnernalin-:V serface the control, systam design with the engine manufacusflr to asbure th~t she operation and performancc defined in prolimantary dcsi~n are attutinable. Accordingly, the continuing relationship btetweea engine and airframe manufacturers is Pratmount during the detail design phase. Detailed diecussion regarding this interface relationship is pro043 vddi hpe ,AC
34
FE
-
USSE
45 EEA 4.GE RA The fuel subsystem con~sists of tanks, refuctiug/ ds-fuehing features fuel feed and vent line-s, fuel purmps vahles fuel gaging comiponents, and Masovialed items such as fuel tank compartment struc3-9
I
.T
subjua.
A typicali fia
is dluatfaed digamp
%)stsms. speal~e1Iy the okns)
pmbimwy i dk *mum$ad fatl sCsA AMCP 70 .DObS o i (W liyu ALT be a Isalka xis . MIL-EA-3636. seqwg as =a~d" by d miss itn waiviy. Tb. (va amberai hW e Its mnbeny ds..~ with Ref. 6 and WOL4TD-230.Dmpa i0s"uoa fasawres reuiw for aauwortistA aV dwarsbed in Chs .3 AMCP W61-A ssmprate and utwrai wsim(oý.g.a. vows maW wrt"t dry, be# areas iew chi&b air~ftd =muams may be opomnd to pmeni igaimios, sousws, sWt be adequatly wfoned haum Are or esplemiom. MIdL-F-J33W also co.th reqaaafwww for the mswrunafsb OanldtsdS sass, whic bs (tad W ausyslff AsU opemuic wA4&We*y. Uwa.. Whosy'm spsdffatwe 3M-ovKdS Iq-
A
mis ptocuuiagI activity prewcilme otherwise. mtc raw'of ambles ass tmpeawi and fuel temperat~~~amt at the miimum wawe amntent of the fuel ABU be a* spciidW inaMIL-F-M383. For asatimumn rcAiibilaty. it is desirable that the coeuation of fue&4dw sabsystems b,, faanuionaliy independent qa the opca~ciun ot o~Ls icWivotiw sub-
'N
AP L
IIA
50
SA A
7A
C01 14
-
APU SNL'IOFF, VALVE iMANUAL),
-
150j~
¶54
ISA
6
8BA -
12A
15
341.1 FeB Tasks Natan tuel tanks normally consist of one or mort bladder-tyiw, tWl intercoauaceed to foami a twn -3 itp"i 1 .t~iy ~~gg e ~ i acz;; -
Ab 10Ab 11AB
1~A IU[IIt
13
df iatm-oth albilxt both um04s. aon She nctwary fuel no% at the pmmramn qedal ey dic cogane model Wimfiara n. Wben posmea. KMw fuel-feed subsystems uh..id bg mocd an Wow as prm monad systems for tacnaad *ta" oft. utabswy In the pararapbsi that follow the pnacqWa &meng twurequcmenas usp-,1acabkc to kq)conmpufant, ast? smmtaons of the fuel stabnystam am daubs wA4 *k,cussed. Reftrain should be ma&ti O hiL-F-fli for additional reqiarmmcnts appisaa a icte Wue maacrials, hardwar. and componcas taed inas a subs,stemn as well as to the samtaflatacte of ths uoa system in the helicoptert The sptcinacaawc. awai dcicribes the dana that must bc pruvadav, to pious the procuring activity to evaluate tbs lue ssabsysaaan design for a new, helcopteu 3U FE USSE O I-N -2FELSSV~M O40ET
BOOST PUMP PRESSURE SWITCH ENGINE-DRIVEN BOOS' PUMP SYSTEM PRiME VALVE kMANUJALý FIREW.ALL QUICK DISCONNE.Ci 12AB -SELECTOR VALVE (MN1NUALI SUPPLY L'NE 13 -HEAlER 14 -. WOBBLE PUMP bABC -CHF CK VAL VE WITH SCREE[N DRAIN VAL\'E (MANUAL) 16AB -SUMP -ENG!NE DRA!NS 17AB
786s
it
F," E EN U
C-L
66r-
SUIPPLY
-FUEL
ELECTRICAL
------
NUlPJE SUPPLIED COMPONENT EI4VELUPL
-----IGINE
NAI
an
1AB -ENGINE DRIVEN FUEL rUtAP 2A;3d FUEL FIL.1 E~ FiL-TER IMF ENDING RN-PASS WARNJN'$jui,-I ~(AIRFRAME MOUJNTED DUAL FILTERS' AiIRFRAME MOUNLED DUAL 'rJEL FILTEK) 4AB 0 DUAL FIrTER BY- ASS VALVE (ELEC*RICAL. 5Ab-
N
I-AL
IIAr 9
hJbdnMoi-
DRAINS
-
)
dance with MIL-T-27422. The type. protection level, and class as defined by the specification shall be appropriate to the application (e.g.. Type I for selfsealing. Type II for nonself-sealing). The tank configuration and installation shall comply with M IL-F38363 and MIL-T-27422. These specifications contain the requirements for liquid-tight structure surrounding self-sealing tanks and the installation of backing boards to protect flexible self-sealing tanks. To prevent overflow each tank shall contain expansion space equal to no less than 3% of the total fuel volume of the tank when the helicopter is in a normal ground attitude. Gravity filler openings shall be so located that all tanks can be filled without overfilling into the expansion space. For pressure refueling systems the level shutoff valve shall prevent filling of the expansion space. Each filler opening cap shall be in accordance with MIL-C-38373. All tanks shall be provided with a low-point drain for fuel sampling and defueling purposes. External fuel can be contained in tanks complying with MIL-T-7378 or MIL-T-18847. The installed location should permit service personnel standing on the ground to inspect visually and service the tanks. Inflight jettisoning should not affect the helicopter adversely. To minimize combat turnaround time, all external tanks should be readily removable and replaceable without helicopter disassembly. 3-4.2.2 Fuel Tank Vents Each fuel tank shall be vented to'the atmosphere :hrough lines whose capabilities are compatible with the performance of the helicopter, without producing tank pressures detrimental to the helicopter structure or to the tank. If a pressure refueling system is required, the venting capacity of each tank also shall be sufficieni .o discharge the maximum rate of fuel flow, without ,xcessive tank pressure in the event that the refueling system shutoff valves fail in an open position. Traps must be avoided, and the subsystem should be operable with the helicopter afloat if amphibious operations are required. If vent valves are used to prevent spillage, they .hall, as required by MIL-STD1290. close when the helicopter is in a position of extreme attitude. The vent lines may be designed to prevent spillage without valves by traversing three directions, etc. See Ref. 6 for vent line design details. 3-4.23 Fuel Gaging In accordance with MIL-F-38363. a fuel-gaging system that meets the requirements of MIL-G-26988 shall be provided. It shall be installed in accordance with MIL-G-7940. System indication of total fuel
quantity and of the quantity in each main tank sli be continuous. Design of the gaging system and fuel cell interface must preclude gaging system puncture of the fuel cell during crash conditions. Each main tank shall contain a device independent of the fuel gaging system to provide a low-fuel warning. The quantity of fuel remaining at the moment of actuation of the low-fuel warning must be suiTicint to allow the engine to operate for 0.5 hr at maximum-range power unless otherwise specified. 342.4 Refuellg ai u The refuel/defuel features for any given fuel subsystem shall be specified by the procuring activity. However. the design criteria for such features shallbe in accordance with MIL-F-38363. All helicopters shall be capable of being refteled on the ground, using a gravity refueling system. without external power being applied. Unless the tar.Ks are too small for the rate to be practical. or the procuring activity has specified another .-ate. the fuel system shall be capable of being refeled, at a continuous rate of 200 gpm without any operations other than removing the filler cap and connecting the fueling nozzle bonding plug being required. During the tank topping portion of the refueling, the flow rate may be less than 200 gpm. If the internal fuel capacity is 600 gal or more. a pressure refueling system shall be required (MIL-F38363). In this case, it shall be possible to refuel all tanks from a single connection, the fuel lines being such as to allow the fuel level in all tanks theoretically to reach the full position simultaneously. On the other hand, the subsystem design must be such that it is possible to fill selectively any individual tank or to avoid filling any given tank. Operation of shutoff valves, plus the precheck system, must not depend upon the use of external power. To prevent excessive surge pressure within the refueling system, shutoff valve closure time may It •r.wtrolled, or appropriate pressure relief valves incorporated. Additional systern requirements are given in MIL-F-38363. Defueling of tanks shall be possible using the pressure refueling system when installed. When the defueling adapter (MIL-A-25896) is not usted also for refueling, positive means, such as a check valve, shall be installed to prevent refueling through the adapter. A typical pressure refueling system is illustrated in Fig. 3-5. A sump shall be provided in that portion of each fuel tank which is lowest when the helicopter is in the normal ground attitude. A drain valve for fuel sampling and for removal of sediment and watersh.ll he provided in each tank sump (MIL-F-38363). 3.11
WARM
czz CA4
10 1-1
cm
1
-J
I
I Ijr'
343S te~-&Addatiouial If in t mWu,6muPing is reqited for rapid reductka of haliepm p grams m. ONe djsIeaa~lpd fuel Ag ot impiWp apwm Mimr*Vst-he hglicopw4. of be diechaqSed into eather a*ares of saaa; ekeritma discharg or Mhe efinwe shaissi plume. Uuln the dumpang. puwps &Wbelh eaved to achive the desired dumping rae Tfhe too %pgM jj capeim snug b tuieu o prevntm find cell collaps whil dumping. 34.2g~ ~ The food system. whavrby fuad is delivered to ahe O.40ims)a sAOl be desiged in acnt wit MILF-39363. to gtamirl. fW moim be available on se anuWaerrupsed barns with"u coatinuous itimatio of the caew. The Weesytms must allow normal terrormanef of dth enginm~z) menet atl attituds normnal for the helicopter on the around, said in both steady and manuve aallaliwds figh p S ad Il~thd~g the servi ce ling. Typically. the fad food system consists of a main h Cm Mihtwf!tns=& relluiareenuts Of' the helicopter within the -points
9ca&incty
*
t.t: zuw t.4UV111611taEU
The fed uiihlpi
'EEC %liut
14W
IZII UllUCt
all required engine and helicopter operating con ditions (altitude and attitude. including maneuvers). Fuel shall be provided to the engine at conditions spacified in the engine specification. The tuction fzed capability shall be dew.rmined for fuel temperatures specified in MIL-F-38363 or as specified by the procuring activity. When the system includes tranbfer taknk(s) aE wtll as main tank(s), the noFMrma squencing of intertank feed must maintain the CG of the fuel system w~thin acceptable limits throughout the range of fuc- loads from full to empty at all required flight conditions. To insure that tht particle 3izes of contaminants in theexeedthe uel o nt imis gien n MI-EW. it maj be necesary for the engine feed syetem(s) to include filters oi strainers. If so, the strainersslwil be in accordance with N4IL-S-8710and the installation shall be in -accordance with M IL.-F38363.
-
indude adquah adaese with NIL-FqId 0ta l i of the syat can bet draeile. AM! madim, bladd taok cavutiu dry bays and pocketa and traps W wsrcuewhm Nt yaA n b rin i tohe stru xirewhreofedw my colect Dain hel orrst a WbeU i.ds w *' Cwp &be tor gd 1AV(a3dUn selwfsealing 2asiks &dbe 0.50 in. d"mowe meisimium Fuel dramn Aab mot be jbsecommvecte with dramn lit carrying otkse liquA The m. Jlation *f We droais Anill be such ttat weda no operating con&tio will drainage re-~te th helicopter or cam in dofgi A
drainag piovintm ina 316.Snfcitdýn
-:o
wbWe brakes. kn those Waisoed waes wben the pamoi00a Iwntho gIC aihf~r OepuwfýPXsiw ea the bioWcannot be avoide. sany joists in Ohw
,mwithin
mialtiengine~~lin hlotesasare. main tiank and feed systunsem be p,-oided for meac engine. These independent systems also must bie so designed that fuel from any tank can be fetto any or all eagines.ThcotosfrtefeSytmAi Twondepnden andisolted ethos ~the provided to move fuel out of each tank. except only one method need be provided for jtttisonable external tanks. Each method ot moving the fuel shall meet
*
neqguemeats and guidelines applicable to the design and installatiwe of feed sysiamnt that will be exposed to awen, $roved rem and for whcK* self sealing tanks ame nquirod also awe given by NIL-IM8)63.
te
occup~ied am esall
shrdoudedan
3.2. ciintres MWd Imafmentom Thcotlsfrheflsyemsllbgopdii egupdn cockpit in a functional manner. A siraplried diagrant of the ruel system sh~llswitch be inscribed oi. the pancl so the functioni each indicated clearly. As a minimum, of presentation ofisthe following performance data for the fuel system shall be provided in
V
; 1
the cockpit:
1. Fuel quantity. c~ch main tank and total 2. Low fuel-level warning roe each main tank 3. Bypass warning for each fuel filter 4. Low fuel-inlet pressure wirning for each engine 5. Indicator lights for electrically operated fuel sltf ae. As additional auxiliary fuel systems are added, appropriate controls and instrumentation shaft! bc provided. 34. TESIING Substanst-itng the capability of the fuel system to tuilisfncon rqir''sdigalphesf aircraft operation is requirWi by MIL.F-38363. Suchl vtrification sAm? occur in three phases: 1. Compn~Ient tcs~inj 2. Fuel sysicem s~mulstor testing 3. Ground and flight testing. 3-13
.
Airflow through the engine compartment is required to prevent the engine: engine-mounted accessories; other components, equipment, or fluids within the compartment; and/or surrounding structure from exceeding max'imum allowable temperature limits, The maximum allowable temperatures normally will be given in the applicable engine or equipment speci-
parilncn! ina. he provided in the applicahle equip;in,-.pecfaicoitior% In the cas of tranmi%,,on, and ecarho.c,. developed bh the helicopter manufat.turer. the heat rejection rate must he calculated. h.icd tn design values for gcar-mesh and bearing efficiencic-. and later confirmed by test (%ec Chapter 4). When the heat rejection rates are known. surface temperature% of individual heat-producing components can he calculated on the basis of free air convection at the ,urface. Rcf. 4 contains a section which treat% each of the fundamental heat transfer mechanisms - i.e.. conduction, convection, and radiation - in considerable detail. The equations and calculation procedures for both steady-state and trinsient heat transfer problems arc given, together with tables and charts of values of the physical properties of material-, needed in the calculations. Should the information be inadequate for a given problem. an cxtcnsivc list of references also is provided. rhe quantity of cooling airflo% required for adequate cooling of the engine compartment also must he determined anal.ticallv by heat-transfer calculatioms. This flow ,,;.ally must be obtained by forccd convection during operation of the engine. with the residual heat remaining at shutdown being dissipated by free convection. The calculation of heat hdance within the compartment is complex. with consideration of all three heat-transfer mechanisms being required. In the design of the cooling subsystem, it is necessary to assure that the airflow over large surfaces such as the engine is such that the temperatures are approximately uniform. Large temperature differentials can result in differential expansion and hence warping of the engine case. Such a condition, which can cause excessive loads on engine bearings and hence premature engine failure. must be avoided. As mentioned in par. 3-2.6.1, an engine exhaust
fication but additional limits may be prescribed by the system specification. Temperatures must be kept below the allowable limits under all operating condi-
ejector ika convenient means for pumping compartment cooling air during operation. The design of an ejector cooling system shall be coordinated with the
tions, both ground 'vnd flight, prescribed for the helicopter for all ambienit air conditions between the hot and cold atmospheres (temperature as functions of attitude) given as limits by the system specification. Further. the maximum compartment or component temperature limits shall not be exceeded following engine shutdown from any operating condition with ambient air conditions anywhere within the prescribed limits. Heat rejection requirements for the engine and its components will be provided by the engine specification. The amounts of heat rejected by other accessories or eouipment installed within the engine com-
design of the engine exhaust system, with care being taken that the installation does not cause excessive power loss or adversely affect engine operation by producing an unacceptably high pressure at the engine exhaust. In any case, the engine exhaust system shall meet the requirements given in par. 3-2.6 (this handbook) and Chapter 8, AMCP 706-201. Procedures for the design of an ejector, or jet pump, are given in Ref. 4. For additional ejector design information see "'Performance of Low Pressure Ratio Ejectors for Engine Nacelle Cooling", AIR 1191, Society of Automotive Engineers. November 1971. Procedures for design, including determination of the
In addition to testing the complete fuel subvstem. all components must be qualified in accordance with MIL-F-8615. and must he so qualified or have pas,.,cd %fetv of flight tests approved by the procuring activity before ground and flight tests are conducted. Fuel subsystem demonstration requirements arm described in Chapter 9. AMCP 706-203.
3-5
LUBRICATION SUBSYSTEM
Engine lubrication subsystems may he an integral part of the engine, thereby eliminating various connections to the airframe and conditions that may lead to oil contamination when changing engines. Lubrication oil may be contained in an engine-mounted oil tank and cooled by a heat exchanger. Engines lacking an integral lubrication subsystem require the addition of an oil reservoir, lines, instrumentation, and a cooler - if an engine heat cxchanger is not available - to cool the oil. Airframemounted oil reservoirs shall not be located in the engine compartment. A lubrication subsystem integral to the engine %will have been tested completely during engine qualification. Component testing of nonintegral subsystems will be necessary in accordance with MIL-O-19838 to substantiate proper rates of oil flow. pressure, temperature. and deaeration. Engine lubrication subsystem demonstration requirements are specified in Chapter 9, AMCP 706-203.
3-6
3-14
COMPARTMENT COOLING
poweir reqavwnamet for aImala. we givnws
Re 4 .
cog. and reobabi1y. SsW wi. of Cgeragration hased
f t =611111 OMW MMawlif Cm~w= suibsyatem must he desumstraW~ by tast. is addiitm to the 011M 1011peratuftld kmsas all VMiiM= CI-
npoi= shape will dqumd up.. helictopr *ac available. weight aNd vokiiimei afe compaable; but coo,. relability. wmlauaaitabihty. and Id cyace coos fey,
0"411uW4 F ANN be WrGWWe from 110u1 Wpi&W. Thais can be awolshe by use of dip he or by cookog the sariams to a asmwalmr balow dUOF. a sysuinm soveriaThe req~rw~ tsefix a ;prip hiom ture survey anc gveii in Chapter1, AMCP 7W]0O3. A systeM temperatture demoowStralies. 1Ipropasio descrnbed in ChaPte 9 of AMCP 1M6-203. may irsquire fmnihu ton"in oddition to the uimperature
The singlesbaft combination bleed APU type. mompream can Moderi siamgie-agW ceaijialS produace 4-0tl and hbghw pressure ratios with a wide range or flow bet wee dhokc arW $,al. Akbotqh the drive compruso may produce higher preames.ra variable inlig guide: vane or diff~ei vanes may be necessary to obtain the required flow range.
3-7
ACCEMSOIES AND ACCESSORY3.2At
rebull in miamiumu Helicopter design rcquiranas ~ ~.~ engie dvc accuoy equwmiets. uxpforthe engine starter and tachometers. mfost acmusories are driven directly from the mawi gearbox. This is to take advantage of the ability of the helicopter to autoroltae in the event of origns failure. If engine failure oc~ ~The dos no fai curs powraccssor ** in istiota of Zeicopter MICA"O" isdiscsse in Chapter 4.
OWER AUXILARY 3-4 AXLAY O E (APUs6)
JNIT NT
34&1 GENERAL The requirement for an APU will be established during pircliminary design. The paragraph describrs design and installation requirements for AMU's and refers to pertinent qualification requirements. Fmrnliasis is nlaced unon the sinale-shaft APU configuration beaue o t ieue u otetedi heicopter design toaard such items as pneumatic main-enigine starting. air conditioning. avionic JR radiation suppression, purging of main engine inlet-protection sys-tems. air supply for antiicing. and the availability of air-driven accessory
I.cooling, Imotors
STU1ODTIL
fit considering the detail design for installation of
DRIVE
(boost pumps. ctc.) additional emphasis is
placed upon the bleed air type of APU. Ncw helicopter designs have favored this type brcause of lower overall system weight, despite the lower energy-transmission efficiency of the pneumrativ main enginestarting system. The bleed air type of APU usually incorporates an intedra! gearbox capable of driving small electrical generators and pumps and, therefore. provides emergency system power of all types. Several APU configurations can be selected to supply pneumratic powei (combined with small amotunlts of shaft power). Four configurations arc compared in Table 3-1 as to geometric shape, weight.
the AMVV all inwefams with the Wir,4pter shell be treated. These include AMt mounting; inlet, exhiusa and bleed air ducting; compartment cooling. arw.4iispor~eAt ussas opitAPsbym.
3823Mtu
1M
l
APt) mounting subsystem xbWl be capabie of withstanding all nlight maneuver forces, providing for eurd). we 9Vc" se.a and beingrwiarrangedsOugvbat. for eas of maintenance (espetcial-
for rapid installation and removal). Driven equipment usually is APU-mounted. thus removing the prublcm of alignment. In some cases. howe.ver. a straight-through ex~ternal drive shaft isA used to transmit shaftf power into an auxiliary or combining gearbox. In such cases, a flexible coupling must be provided. The APU often can b- supported rigidly. This is done conveniently by means of a three-point support arrangement. Two pin-type mounts on each side of the APU limit vertical. axial. and horizontal motion. !3ut provide lateral freedom for thermal expansion. A singlc gimbal then is used to suppoirt vertical loads.
while giving thermal freedom radially and axially.
In most helicopter applications, vibration isolation has not been required because the APU is able to
withstand flight loads and vibrations without shock mounts. Exprience has shown that frequencies up to 500 Hz are significant. but that APU susceptibility is highest from 5 to 100 Hz. The APU must be capable of withstanding the complete aircraft vibration spec,truni in both operating and nonoperating modes. When the APU is not operating, normal loads are not available to stabilize parts in piosition. Hcnce, it is possible for external vibral'on to cause unusual motion vf iniernal parts, thus resulting in excessive wear and premature failures. APU comnponents of parlicular inierest In this regard are the combostor liner assemblies, the acrodynamically located internal 3-I5
Al 14
,
TABLE 3-1 ArIU TYPES FOR MAIN ENGINE STARTING ENVIRONMENTAl. CONTROL. AND ELECTRICAL SUPPLY
I INL-HF DRIVEN
SI NGLE-SHAFT
SINOE*SAFT SINGE-SAFT
SN.SHF
COMJlNATION BLEED
JCOMPRESSOR
T1
COST
100% 100% 100%ý 1000%
REI -IARIILITY
100%
LENGTH DIAMETER WEIGHT
*
COMBINATION
BLEED-ITORQUE CONVERTER
1h
--- iC--T
(710% 850% 970 125%0 91",
-
1
TWO-SHJAFT
PT DRIVEN COMPRESSOR
C
1G10%0 000% 10061 125%o 70nG
TP
84% 00%1
-
16000
73%1
ATS-AIR TURBINE STARTER C-COIMPR ESSOR PT- POWER TURBINE T-TURBINE 'Vj-ALTERNATOR ECS- ENVIRONMENTAL. CONTROL SYSTEM S-STARTER of high-temperature exhau:;t ducting, even with some seals, the gear train, and the nonpreloaded bearing sacrifice in inlet or bleed duct lengths. asemblies. Bccause some installations do not require .rborne APU operation. considerable time ma) be ai 3-.2IneDutg accumnulated in this nonoperating mode. An82. inletaiDullctori ufgfe i eige n A ne i olco rmf fe sdsge n ~~~~in addition to the requirements of MlL-P-8,586. the operauing and nonoperating modes. These ,.h//I be established based upon the vibration spectrum at the APU mounting points and not merely upon the freqeiisof the main rotors. The :rend in maintenance philcsopN)y is tossard minimum scheduled maintenance of the APU., but with the desigim adapted for rapid APU removal in case of malfunction. Hence. quick-removal connections shwi! be used. with airframei componenits and support structures arranged so that the unit can be removed without removal of other equipment. Minimum system weight usually will dictate that the APUJ should be located near the main cr-gincs. A multienigine aircraft, with air ducting required to %tart each engine, suggests a submerged APU position for minimum weight. Pod or surface locations can be used when fuselage space is c-itically limited. Also to be considered is the advantage cf a minimum length 3-16
ducting then runs between the fuselage surface and th,, collector connection. Ducting pressure losses will iiffect API) available power dircct!y. A typical relationship is illustrated in Fig. 3-6. For any output shaft power level. a correction factor is given showing the power loss for each unit of pressure loss in the airframe ducting. Values are given for both inlet and ex-
haust losses. The APU manufacturer may prescribe limit%, on these losses in a form suz.h as is shown in Fig. 3-7. If the sum of inlet and txhaust losses falls btvlou~ the curve, the ;equirements have been met. The APLJ air collcctor design may be critical. especidlly %hen transsonic compressors are used. Entry conditions to the inducer will have inmportant effects tin comnpressor efficiency. Collector design can be compromised b) the location of accessories, Viut an attoimpt should be made to mnaintain at uniform total pressure distribution around the csompressor inlet.
9
AMCP 706-202
FUEL FLOW PARAMETER CORRECTION:
Wfi
WbCORR. WHERE Wf/b CORR.
lb/hr
FUEL FLOW PARAMETER CORRECTED FOR INLET LOSS, lb/hr
Wffl5
FUEL. FLO'.,V PARAMETER WITH ZERO LOSSES, lb/1hr AMBIENT PRESSURE, psia PRESSURE LOSS, in. H2 0
Pa
AP INLET S
(S P8 ZPINLET)
=INLET
~~
Pa14.7
0.8
-J
W 0.6
bEXHAUIS1 LOSS-
I
0.4__
__
0.2 0
10 20 30 40 50 60 70 80 90 100 OUTPUT POWER PARAMETER WITH ZERO INLET & EXHAUST LOSSES, hp/6, lhp
110
Figure 3-6. Fi'ufmnuace Corrczioas for Dvel Lowsa inlet. The airframe ducting should include a straight
I
I 0.6
I
tstratification and cavitation in the ductinR.. If a device is incorpo~atcd to protect the inlet against dust or foreign object damage (FOD), this ~may affectthe velocity profile and the resultant rs sure losses must he coordinated with the APU manusothat mdlspecification performance is maintained. One technique is to use directions.1
Ill
0.,facturer
a--. . rAFIAUST
changes of the total duct system for separation of intake screen shall be used for FOE) pro-
t IAid
0.2_ .
0.G 0.8 uc r PRESSURE Lu5S, P5.
1.0
~larger particles (>200 micn-oits). 1.2
Figure 3-7. Allowable Combined Inlet and Exhaust Duct Piressre Losse stabilizcd and directed by splitters or baffles. Thc APU will have channels to provide contit oiled accelecation of air from thc collector int3 the inducer
<'possibly
tectioi,. Tisi -nay be located at fuselage entry. AIPU collector entry, or compressor entity. The discussion of the cflgitte air induction subsystem in par. 3-2.5 is eial plcb- oteAUilt 3-.2.2 Exhaust Ducting Exhaust duiting design will be related t,, APUJ compartment cooling arrangements. The gaii-exii ye3-17
AMCP 706-202 locity can be used as the primary jet for an ejector. Stcondary air is taken from the APU compartment. Also, the ducting is a substantial heat source and should be located to minimize effects on compart-
Some components arc APU-rnounted and othcrs are airframe-mounted. Electrical power to activate the system may be provided by the aircraft system or by an APU-driven generator. APU controls shall be
ment temperature. If compartment ventilation must be limited due to fire: hazards, high-temperature bellows can be used to seal ducting. radiation suppressionand mayduct be needed for theAlso APUIRexhaust. Compartment cooling methods used on the main engines will apply equally hete and aro discussed in pars. 3-2.6 and 3-6.
qualified concurrently with qualification of the APU in accordance with MIL-P-8686. most S advanced d anedAg APU sequencing s "ihe most is done with
3-8.2.4 APU Bleed Air Dwting In the case of a bleed air type of APU, a collc(tor arrangement will be integral with the unit. The APU control valve can be mounted directly on the APU bleed collector flangc. If installation ,pace limitations cxist, or valve weight will overload the APU flange. the valve may be line-mounted. A flexible bellows can be used to minimize flange loads, thus accommodating vibration, thermal 'owth. and installation tolerances. A quickdisconnec' type of connection is desirable for maintainability. Attention must be paid to duct pressure
solid-state electrical equipment. A speed signal is obtained from the gearbox or power section by mcanm of a frequency signal from a magnetic pickup, tachometer generator, output alternator, or mechanical spced switch. A frequency-sensitive sequencer then actuates relays for slarting and APU accclcration. The start is initiated by a switch that actuates the APU starting system, This may be electrical or hydraulic (par. 3-8.3.5). The APU will begin rotation without fuel or ignition to provide a momentary air purge of the air and gas passages. Fuel and ignition subsystems must be actuated at the lowest possible speed, perhaps 5% or below, to insure good starts at both cold and hot extremes of ambient temperature.
loads. APU model specification perfovmance is based on pressure, flow. and temperature a' the APU bleed collector flange. 3-8.2.5 Cooling Surface temperature limits will be specified in the model specification. During APU .)pcration, the compartment cooling system shall cool the compartmcnt adequately. Temperature transients should not exceed the limits specified in the model specification.
Actuation can be accomplirhed by a fuel pressure switth, a timc dclay relay, or I speed signal. Some combustion systems require separate start and main (or run) fuel systems. At 10-20% speed, the main fuel valve will be opened. A third speed point can be used to turn off the starter. At approximately 90% speed. start fuel and ignition will be turntA off. This signal, with a time delay relay, also can arm the aircraft load circuits. A fifth sequencing point at 110% speed pro-
APU firewalls are identical to the main engine requirements discussed in par. 3-2.4. The reduction drive, accessories, and lubrication y.en. reprent . onsiderable heat. sore at Pon be dissipated into compressor inlet air or into the APU compaitment. An oil (to air) cooler may be needed to maintain system oil temperatures within acceptaole limits. A typical small bleed air APU may reject about 150 Btu/m:n to the lubrication system at sea level prossurc and 130'F. An exhaust ejector could provide a cooling airflow rate of about 5 lb/ min. Compartment cooling is discussed further in par. 3-6. 3-4.3 APU SUBSYSTEMS
vides for protective shutdown.
mrnasniach as APU) .ubsystems ai similar to those previously discu•ad in this chapter for the main nigine(s), the paragraphs that follow discuss only those characteristics peculiar to the APU. 3-8.3,1 Electrical Controla APU elec•ricisi coatiols are categorized as se. quencing. protectivc, and toad, or output, controls. 3.-18
38.3.:.2 Protective Controls Protective controls arc required to confine malfunctions to the APU and, thereby, to protr.t the helicopter. Protective devices may include overspeed, exhaust overtcmpcrature, and compartment ovcrtcmperature subsystems. A simple thermocouple sensor in the APU tailpipe, feeding to the solid-state circuit control (sequencer), provides APU overtemperature protection. Similarly. thermocouples within the compartment can signal the sequencer as fire protection. Built-in test equipment should be used only to the extent necessary to indicate an APU failure. Indicating lights shall be used for pilot advisory purposes. Audio annunciators can be adapted for the if desired. APsu 3--.3.1-3 Output Controls The elect-' -1 load-control circuit usually is armed by the 90% ed sequencing point and a time delay
I...
AMCP 706-202
relay. Acceleration israpid from 90 to IOD% speed for
governing throughout the operating envelope dcrined
APU, hoeemyhv mcucteaclrto
Techniques for scheduling fuel may bc mechanic i, ithe oniode Th lcrncsystemis at-
yasfe ;-cbten9
av n
o 0%
contol beed vlveto ir a argulae uncton o ex haust gas temperature. A thermocouple in the APU tailpipec can signal a solid~statc circuit componert, which, in turn, gives a modulating signal to the load.control valve. I he valve will open so that maximum continuous exhaust gas temperature is maintained. Thus, maximum bleed air available from the engine is obtained. Large shaft power requirements require careful examination of transient opt.-ation, because a rupid response may be needed to avoid overtemperature shutdown. If the compressor is marginal on stability. the load control valve may be s'heduled to bleed small amounts of air during transients to prevent compressor stall. Output load controls for electrical, hydraulic, or direct shaft power arc regulated by aircraft system
Icomponents.
3-8.3.1.4 Ilectirk-al Control Location APU-mouznted control components include all driven enuinment: sensors for temperature. Dressure, ads!%ccd. valves and ignition components.-anti other control components capable of withstanding ccmpartment temperatures. Solid-state sequencing controls. p~ower suppily, load-contsol valve controls, miniaturized relays, and malfunction in~dicators shall be airframe-mi tinted so as to lai-it temperature; to 2001I or below. If required, some of these components can be mounted on 'he air inict collector oi other ducting %vheretemperatures cani be limited by heat transfer to incoming air. M3-.31.5 Elecirk-all Prower Reqsareme ts If the aircraft can provide small amiounts ofeclectrical power, the APU system is simplified. About 4 A, 24 V'is adequate to operate most APU relay s3-stemb,. Battery' systetms suffer from ambient temperature limitations in that the battery must be kept warm (00 F or above) to permit the -65OF APU starts required by MIL-P-8686. If no batteries art available, an APU-drivcn ignition and control generator can supply start and run sequencing power. Such a system requires a storedenergy APU start, such as hydraulic or pneumatic. Voltage buildup must be very rapid so that system sequencing can begin early (5%) to achieve low self-sus~\ taining speed and good cold starting. ")3.3.2 Fue,! Sysiein Controls The APU fuel system ý;ontrol% must provide for automatic starting. acceleration, and rated speed
orpecidicatr
sped cotro. prtectve ircuts retcrequirequrdd In most cases, however, the APU system requirements are simple an~d are handled best by a niechanical governor with aceic.ledtion fulnwshdldby compressor discharge vre' aure. The APU fuel system cnitofhefuel supply (common with the main engine(s) ), airframemrounted boost pump, and fuel lines (with shutoff and check valves) connected to the APU. An inlet filter of large capacity, but small micron rating, provides clern fuel at the APU fuel pump. 3-8.3.2.1 R~ted Speed Govevning The APU speed-regulation requirements genemitly are satisfied by droop governing, as explained for the main enginc in Chapter 8, AMCP 706-201. A speed band (droop) of' 2-4% through the load range is adequate for frequency coitrol at 400 Hz AC power. Peovided bleed airflow and pressure requirements are ment, the speed band is not critical for a bleed air .ArU. Sopccd recovery and stability %villbeespecified in the APU model specification. It reqi'ired, isochronous gover ning to hold speed within a narrow bond (±0.25%) can be acccinplished by a null system. Topping speed adjustments should be provided to account for installation differences and deterioration of the efficiency of the APU or driven equipment between overhauls. 3,3.2FlrngRjrmes 3-...
Fitrn
eqieet
Filtering has been a problem on some military helicomter installations, Combat situations have resulted in .fuel contamination beyond original expectations. Servicing oi components in the field also can introduce contami nation, which must be considered. To keep required maintenance at a minimum. AI'U Filters should be of extr..-large capacity, or of self-purging configurationi. A high percentage of fuel control, valve, ance nouzie failures results from contaillination damage. Sources of con~tamination in-
vlude line or component contamination duriing servicing and wear products generated interiially in pumps and m.)ving parts. Screens or inicronic filters shall he placed a~t va ,c aod nozzle enti ances to prevent passage ot pa! ticles left during assembly, installation, or field maintenance. A PU fuel rump midi and ou- let fiPers.'Jiall he stand;ard equipnicnt and of the throwauay type. The rated micron size of filters should he as large as pussi3-:9Q
,
AMCP 706-202
*
ble, considering orifices and jet and internal tolerances, so that capacity requirements can be reduced. 3-8.3.3 APU Lubrication Subsystem The lubrication subsystem generally is selfcontained within the APU, unless an external oil cooler is required. The trend is toward a completely sealed oil system requiring no scheduled mairnenance. Filters are included internally and, with current uiits, are serviced at specified intervals. With adequate seals and filtering of buffer air (if used), external contamination virtually can be eliminated, Slightly larger filters, with bypass valves, should eliminate the need for change between over.hauls. Oil consumption rates generally are low enough so that oil level.checks may be eliminated for long periods. As with fuel subsystems, field maintenance of APU lubrication systems introduces more problems than it solves. Reliability factors for lubrication pumps, filters, relief valves, and jet or mist supply systems are high. Altitude operating requirements are def--ied in the APU model specification. Maximum temperature limits also are specified in the APU model specification. Qualification testing of the lubrication system is concurrent with APU qualification in accordance with MIL-P-8686. 3-83.4 APU Reduction Drive The APU-reduction-drive design is established primarily by the driven equipment and accessories recuired. The bleed air APU usually is designed to deliver a small percentage of its total output shaft horsepower. Drive pads will be provided for fuel controls, control generators (if needed), and all driven accessories such as electrical generators and hydraul"ic pumps. ic pumps. "The turbine nozzle and diffuser matching within the APU can be changed to trade off shaft power capability for bleed performance within the limits of stall. For: the shaft horsepower APU, output may be concerntrated at one or two larger pads. If APU power is applied directly into an aircraft combining gearbox, fewer APU pads will be needed. In this case, required helicopter accessory outputs can be obtained from the auxiliary gearbox. APU Sts 'tilg The APU starting subsystem shall be fully automatic, using the sequencing systems previously described. The starter energy level must exceed the APU drag (resistive) torque by sufficient margin to provide the required acceleration. The starting torque requirements will be described in the APU model specification.
3-8.3.5
3-20
(. Electrical starting is satisfactory if starter and battery energy levels are chosen properly. The minimum-weight subsystem requires the smallest battery consistent with adequate breakaway torque, including consideration of initial voltage drop. This will result in a longir starting time to rated speed (15 to 20 sec), but battery is adequate to provide some torque to very high speeds (80-90%). The basic limitation is that cold-day starting (-65°F) is not practicable without warm and oversized batteries. It will be found that when APU size increases above 300 hp, battery size becomes excessive. Also, the slower start makes the APU more sensitive to the fuel acceleration schedule. Achieving successful atarts at both -65*F and +130°F without adjustments may require a compensation mechanism. The hydraulic method of APU starting is satisfactory, especa~iy when -650F starting is required. Hydraulic starter motors should be sized for high initial torque to give rapid acceleration. This will tend to result in a lower accumulator volume requirement, and to reduce sensitivity to fuel schedule variations. A typical subsystem will have an initial cranting torque of 50% more than the highest APU resistive torque. Ideally, the motor should incorporate an 'verrunning clutch so that no drag is induced onto \: the APU when accumulator fluid is expended. During -65°F starting, the APU may self-sustein at 40-50% speed. A few foot-pounds of drag from a motor can necessitate additional accumulator volume so that starting torque continues to 55 or 60% speed. As previously discussed, hydraulic starting can be arranged by using a battery for control power or by incorporation of an APU-driven ignition and control (permanent magnet) generator. 3-8A4
RELIABILITY Rlait R haraciLiTe
"
Reliability characteristics are specified in the A'PU model specification. For helicopters using thc U for inflight emergency power, starting reliability is o major importance. Typical requirements for APU starting failure rates are shown in Table 3-2. TABLE 3-2 APU RELIABILITY _
_
NUMBER OF STARTS ALLOWABLE FAI LURES 0-500 501-775 776-1050 1051-1390
H
0 124 3-
i
-,
.9AMCP
APU operatini life may be dcmronstratcd to :kwy given rmodcl spech ication rcquiremnent. A test plan may be chosen from~ MIL-SI 101-7811. Typically, if a mean time bctwecn Failures (NITBF) of 1500 hr ii, specified. a test with~ sevcn APU'r. each operated to 300 hr without failurc, would bc recoimndvcrced Othei combindtions of run time and number ol e ngintes may be 4:hosen. e.g.. three APU's with each iunning 750 hr without a relevant failee. High installed reliability can be obtalnt-d by spci fying a time between overhauls (TBO) close to the specified MTBF. However, this rvutts in a considcrably higher life-cycle operating cosi. Lc west Cost is obtained by u-.ing :hc rmiove for ailiure (RFF) phikctophy, whck. APU~s are repaired or rirmoved only in event of a malfuntction. A fail urr mode and el'i'ct analysis (FM EA) for the APU shall be specified to show tftc consequences ol each ;,tobat6.c Ivilurr. This will assist in reliability prediction, choice of scheduled inaintenanct inter-
*
..
B fRFvru vlospis.adtedtriito losopies.rotor iecral problem areas have been expericirz.:d by oprnio of the APU in the heclicopter. These &amaa incaude vibration, irlcircualton 0f exlu~ta gsrz > FOD, eoý. They point up iVhe importance of careful delineation, in the quality assuranc provai~is1 of the A PU model 4ppv-ification, oý d.-Aign and tesi ronclitions that simulate the helicopier eflvironmiria. The vibration environment is primiary. The profile
of amplitude and froquencies at the APU installation must be defined. The amnplitude and ficquencies of the AlFU-gencratcd vibraiion also should be spc~ifled to anticipate airframec structural probleinis, APU air inlet iemptrature limits arc specified, and the APU installatioan ust be desigtned to prevent recirculttion of main tngiiac or A?t3 *txhau-ýt gas. Carteful aocntion must be 7.iven to thi.t prokilena because ai is impractical ii' attemrpt to limit heiicop ter operfatiori in undesirable wind directions. Ir.1ut duct locatioav. either must be rtmote fianm exhaust outlets, or sdfety shutdowii sen~oms must be pro-
vided. Similar arguments apply co~cerning FOD. stird, or dust inglestioin. Genetally, higi inlet doct locat iors are pi cferable bec-ause concen!vtion i. dust in huvem is straxified vertically. IVany APU strtic-. problems ca!. be traced to fuel aysteen components. St.mc of these re- 61t ffrom helicopter fied system contamination. Addit~onal trrphasis should be giv.-n to filir~tion, and to Pk'C. oenur of main ýRnk contamination by proper fuelhandlicip methods. A major sou ce :)f reliability problems arise& fromi
-'
706-202 maintenan;ce reqluireenirts cith.cr scheduled or unschcduled. For example, more contamination is in. troduccd into oil systcms (causing exccisive wcar and early bearing failure) through frequent oil level checks, oil addilions, and oil changes than through seals arid vents during normal running. APU design should stress minimum scheduled riainitcinacie, throwaway filters and components, scaled systems. and automatic controls requiring no adjustment. This approach not only will increase reliability, but also will decrease not life-cycle AF~U costs. M,5 SAFIETY PROVISIONS Good APU safety design must inciude provisions to prevent a failure frorn causing helicopter diemage, and. if possible, to permit mission completion in event of a failure. Thus. the APU installiption shall be designed so that fire, APU rotor failure, and crash damage arc contatined within the APU compartment. APUJ-rotor constainment is an important safely cnieain n a ehnldi eea as Strulturc can be designed to withstand and hole a tni-hub burst at overspeed trip condlition4 (a.ssurniing a fuel control failure). but this causes an une~e'irablc weight penalty. Alternatively, rotor integrity car, bc. d-mcnstrated by in.....J.p.c-.-.-lizc tests to micarure stress !evels under operating coitditions. Furtkarr. systems cani be arranged to gv'Arantee that blade failures oct.-ir first (e.g., stress groovc.9, but that smaller mass bladr fisiltirc cart be contained within the casing structure.
Fuel and ignition sour-ces s1haý' bc: separated by m~eans of the compartment deiign. On'ý philosophy is to put the entire A PU into a fireproof ci~nip-irtmcnl. A seooaid philosophy sodi-s to prevent f1ie by confiniing 1w.) sourucs and by s.-gii.gating the hot scction %v'itha bulkhead. The APU coritrv~As arnd oil sumn should b,: housvo in fireproof contair'crs. Electrical itibsy-terr ignition sourcce., should be roijted or housed away hornm bel lines. APU inlet air sh~ovld be dlucied from outs'd. thc helicopter to prevetit rocirctrlation in case ol cornuartinent fire. [ire detectors and fire wxinguishing cquipwmemt shall be. u~vd to pr'olcc: against firc within the compartioien (see par. 32.4ý , ll APU fuel system romponents shall te crashworthy. Fuel line-s shall be made o: flexible hose with steel-'hraited outer fheath, with the mtn.'rum ni-niber of .ouplinps. At bulkheads, the hosce should bc run through unc'it, using fr*mngibic hose stabilizer fittings. Whcn lines So through w firewall, selfscaling, breakaway couplings shall be used. All line supports should be frangible. Lines shalt' be 2CLI30% lo~igce. than necessary to aicomroadate structural dispiacements.. Routing shall be aong the heavy basic 3-21
AMCP 706-202
-tPreferred
*
structure. but away front electrical comp,.mnents (unless electrical systems art shrouded). D~rain 'incs for -.onbustor. fuel pump. gearbox, vents. etc.. shall be connected with frangible fasicrners. and made of low-strength matcrials. design calls for enigine-mounted fuel bogst pumps with suction fuel supply. In the event that tank-mounted boost pumps are required duc to fuel subsystem configuration, they are acceptable if meunted with frangible attachments. Electrical lead wires must be 20-30% longer than necessary. and shrouded to mirimize crash danmage. Self-scalintE, bicakaway couplings shall be used at all c )nnections. Filters and valves shall withstand 30g aoads applied iii any direction. Electrically actuated valves can-hec bulkhead-mounted, with wiring on one side and the valve and fuel lines on the opposite side. APU oil tanks and coolers also shall withstand 30g loads. and 'must be mounted away fte-icn impact areas. They shall be located within ttue compartment. but away from hot sections and inlet air ducting, to prevent ingestion of spilled oil. Oil filters shall be integial with the APU. Ratteries and clecoical accessories shall be located high enough in the fusciage to remcve thanm frt'ii possib:, fluid spill~gc areaý. They vA~yll ttt comnrtrt-
3-22
mntcnaliscd with lnexible fire resistarnt pancls. Extra %."rc length is needed and s~hall be supported; wilh frangible connections. The basic structure .thall %ith. stand 30)-g loads applied in isny direction. REFERENCES I . C. R. Bryan and F. F. Fleming. Some InternalFlow Characteristics ci Several Axisvnimetricol NA CA I-Series Nose Air buttis at Zero Flight Speed. NACA RML54E 19A. July 1954. 2. K~uchemtaiin and Weber. Aer-odynamnics of Propulsiion. McGraw-Hill Book Co.. NY. 1960. 3. J. Seddon,. Air Intakes Ja~r Aircraft Gas Turbines. Journal Report. Royal Acronaoticall Society. October 1952. 4. SA E Aerospace Appnlied Thzrrnodynamics Afan*a'l. Society of Automnotiv. Enginccrs. NY. October 1969. 5. T. Himka and R. D. Sekmple. Enginc/Tranx.mission/.4irframne Advanced Jn~rg ration Technique~s. TR 75-16, USAANIRDL, Fort Eustis. VA. May 1975. 6. N. 0. Johnson. Crash worthy Fuel System Design Cr-iteria and Analyvsis, TR 71-8. USAA I'Al DL. Fort Eustis, VA, March 1971.
AMCPX70202
CHAPTER 4
TRANSMISSION AND DIIVE SUBSYSTEM LI4T OF SYMBOLS A
C -capacity of bearing far lire of ii! cycles with 90% probability of survival..' - case convection cooling cotfricient, AP0 - load rinclnation factor (helical get), dimensionlest
- liner (steel)O01, in. speCuic reat of oil, iiitul ln-t F1 b geSar diameter, in. d D actor - macrial buring), dimnsionlss D~c- bolt circle diameter, in.Dp)- stud pitch dianieiter, in. 0,, - anvolute base circle diameter, in. DI - inside diameter, in. D~in - spline minor diameter, in. D, - ouarsroo diameter, in. DP - oitsch diamecter. in. D, - pitchoodiameter, in. D -=Major diametfer of spline, in. D2 - outside diameter of spline tooth mnember, in. d - pinion pitch diametcr, in. d - track of breked wheels, ft d -' light alloy &Wcion OD, in. E - modulus of elasticity (Young's modilus), c
e C_
£
)
- frequency of backward tiavelling wave, Hz race curvature, %of ball diameterr
baxing fo ccolu-iner
B1)- life at which 10% of a bearing population fail, cy,.Ies or hr b Hertzian contact band semiwidth, in. b -bea~ing 013, in. Cc. CA
fa
psi
processing factor (bearings), dimensionless ED energy dissipation rate, Btu/in.-min El combined modulus of elasticity, psi r -pitch plane misalignment, in./in. F - flow. rate, gpm F - face width of gear tooth. in. F - lubricztion factor (hearing%). dimensionless FA, -bicakaway blip for"e. lb F, = effective face width, in. F., - average effectivC face width, in. f - coefficient of friction, dimcniionicss -
DESIGN
G
EHD matcrial paramecter, dimensionless - speed effects factor (bearings), dimensionIns - lengthwist tooth stiffness constant, pjs
h hE
misalignment, factor (bearings), dimensionloss ftoil film thickness, psin. - EHD oil film thickness. #sin.
G G (is H Ii
-
1
specific weight of oil, lb/gal
7 ý
of inertia, slug-ft' Imoment i geometric shamc factor, dimensioniexsst-l
K K K, Km, K, K,
K KI
K, k k k L L
-Hertz stress index, ftpsiniols -strcss concentrationfatrdiesols nrifatdmesols - life factor, dimicnsionl~ss - misalignmricrt factor, dimensionless - overload factor, dimensionless - reliability factor, dimensionless - dysicze factor, a dimensionless Mtemertue factor, dimensionless -tenmperlatur factor, dinensionless - conversion constant - con~tact line in~lination factor, dimnensionless - geom~etry factor, dimensionless - gear face width, in. - design life or scheduled removal time (TBO), hr
LA
=
Lrj L2
-
L10
-
M Mf Inl Mir,
on, Opt"
i
-,-
Adjusted life, hr
gear center distance, in.
- life for 2% failure of a bearing popuiatlon.
hr life for 10% failure of a bcaring population, hr -mechanical advantage, dimensionless - moment, in.-b - prortle contact ratio, dimensonless
.
V Y
= gear rutio. dimensionless
=Con~itc! ratio factor. dimensionless modified contact ratio (spir~l beve; g5or), dimensionless
-
/*
4-1
number of teeth (gear or spline) or bolts or sttds
N
TA
- ambient air temperature (cold condition)
-t number orteeth on gear
T,
- critical temperaturm, F
- number or teeth on piniot
7".
-
T s
- circular tooth thickness, in. - initial temperature, °F
N
- number
NI NP
n n RN nap
PL PLC Pc P ?d P4
number of discrete values number or radial nodes - rotational speed, rpm - critical speed, rpm normal operating rotational speed, rpm - pinion rotational speed, rpm - load, lb power loss, hp - power lors to oil cooler, hp -power loss to oil cooler (cold conditiwi), hp
Pf P, P, P,
base pitch, in. - diamet'al pitch, ir.' - transverse diametral pitch (measured at large end of bevel gear), in.-' - friction power loss, hp - pt ver inuut to transmission, hp - power loss, % of transmitted - mean transverse diametral pitch (bevel
P, P, P p
-
T. U VI V2 VT VT V, W W
w Y Y
-
Y4
W Wd W,
Sgear), in.-
oil pump los,, hp fastuner tension !oading. lb gear windage power loss, hp pump discherge pressure, psig
- average external surface temperature, "F - average external surface temperature (cold condition), OF -- EH I speed parameter, dimensionless - rolling velocity of faster of two bodies in contact, it./sec or fps - rolling velocity of slower of two bodies in contact, in./sec or fUe. - total rolling velocity (V, + V2) of two bodies in contact, ln./wec or fps (Vi + V2)/2. in./sec or fps - slig "4 vldoc;ty, in,/sec or fps - load, lb - helicopter weight carried on braked wheele, lb - EHD load parameter, dimensionless - dynamic load, lb - failure load, lb gvat touih luad, lb -- effective gear tooth load (K. W, + W).), lb - running load, lb/in. - modified Lewis form factor (spur gear), di-
-
SPM
OF
W
-
circular pitch, in. -
- modified Lewis form factor (helical gear), dimensionless
Qs Q
torque, lb-ft or lb-in. brake torque. lb-ft skid torque, lb-ft - flange torque capacity. lb-in.
2
- modified Lewis form factor (bevel gear). dimensionless - total transverse length of line of action, in.
0.
-
stud torque. lb-in.
7
=
R R
-
mean transverse pitch radius, in. reliability (for I-hr mission), dimensionless
Z,'
-
R,
-
distance from pitch circle to point of load
a
- pressure viscosity coefficient, in3/lb
a
r r
application, in. - radius of curvature of g-,ar tooth, in. -radius of curvature of pinion tooth, in.
- linear coefficient of thermal expansion. in./in.-*F - contact anghe, deg
-
.nrcbability of survival, dimensionless
S S.,
rms surface finish. pin. allowable endurance limit stresi., psi
SA,
-
S, S,.f S S4 •S, S, S,, T
- compressive (Hertz) stress, psi - compressive (Hertz) stress at failure, psi - hoop stress, psi - bursting stress, psi shear stress, psi - tensile stress, psi - torsional shear stress, psi - tempcrature, °F
TI
-
4-2
bearing stress, psi
ambient air temiperature, °F
4
mensionless
normal circular pitch (helical gear), in.
i
_11
.
.- ,,,,, a
rn fam~ ,.,,
,e,,onsn,.,
modified scoring geometry factor kspur gear), dimensionless
- fraction of theoreticai contact (splines), di-
mensionless - incraent, as A I, deg F or A(D,,/ 2 ), in. 6
-
&• a 4
-
mean CLA surface roughness, pin.
incremental lrowth, in./(in.-*F) - deflection, compressioo, or protrusion, in. - efficiency, dimensionless or %.subscripts c ana f, p. and i denote coarse %'idfine pitch, pump, and transmission, respectively - ratio of oil film thickness to surface roughness, dimensionless -
failure rate, hr-'
C
.
a)dynsric viscosity,Ib.-sc/in.1
FS
made lighter, more efficicnt, art d ss costly if it werc
Poino)-is ratio, dimensionless v ..heict] tooth kcad line inclination, deg EHO parpmeter, dintcnsionleus 3 -denshy. sltrg/ ftl or lb/in. a -standard deviation, dimensionless gear tooth pressure aingle, deg - ormal pressure angle (helical gear), deg 0, - ransverse operating pressure angle, deg 0 - gear tooth helix or spiral angle, deg - angular acceleration or devele.-ation cf rotor, rad/icc' W rotational speed, Hz
41INTRODUCTFION 4-.1 GENERAL Thc proper use of this chopter as an aid in the
achievement of satisfactoiy transmission and drive
syscmdetildesgnreqirs aclarunderstanding of asi Te cntcitsrefectthe past in coceps. stvcal generae anhe suggdested analfor tio u-nrdw..rhac sinr te~chnanurs
9
not necesary to conrider interfaut effects. Hiowevecr. the design optimization technique-,. addresscd in this chapter are limited to the components of the powvcr transmission subsystrms without consideration of i be possible ovcrridirig effects peculirar to a given aircraft configuration. The designer must be aware of the significanice of thc to*.al Army environment. Major subsystem comnponeiits such as gearboxes, driveshafts. or hanger bearing assemblies probably will be subjected L,) rough treatment daring shipping, handling, arid removal from or installation on the helicopter. The consistent use of sophisticated or special tools and torque wrenches simply will not occur, even though specified by the designer. Extremes in temperature. humidity, sunli-ght, precipitation, and sand and dirt cotmainwllcurdinbohfgtopac :onamnaion wqill ocur urig bth nigh:topera. tion and lengthy periods of outdoor parking. Pres. tawtr sr laigcupct-crilyn
mal shock. Exposure to such hostile envionments will cu eetdyfrln eid ftmbtms not compromise the mission availability of the hhelecpter. Improper maintenance, tool drops, and rvreo mrprisalto hr osbeas willrs o r. Compro nernstaldesion wustre toslern ando
desiner ustpracicehis kil fro a ovin or The use of geared transmission systems predates irecorded history. A relatively sophisticaled differential gear-drive system was employed in the Chinese
forgiving of such treatment wherever practicable.r
*C.Ln...&I
~
~ ~D.C 10
%An
i
and geared drives still represent the most efficient mthod of power transmission. This chapter is intended to encourage rather than to subvert new and unconventional approaches to old problems. The sole limitation upon incorporation of the unconventional in helikopter drive systems is that the ba:sic rules of nature (la~ws of mechanical physics) are relatively inviolable and should be treated w'ith respect, There is no perfect or unique design solution to a given power transm~ission requiremenrt, and all known designs have been compromised by the individual requirements of the aircraft into which they must be integratid. Optimization of a design cannot e viewed within the context of the power tiansmisints alone; thorough trade-off studies must such factors as suspension, layotit. airframe ,]support structure, rctor control systems, aircraft weight[ and balanc, space limitations, and locations of eniginecs. Almost any known drive system could be
[
\consideir
K'.-.
helicopter and its drive subsystem; thcs:: cleaning solutions may be more than 100'F hotter or colder
tebasic selections of gear, bearing, and shafting repiresent the present state-of-the-art, leve, te tehnoogyas dscused oesnot exclude the future in that areas of uncertainty and limitations of knowledge are emphasized wherever they appear applicable. The mechanical drive system 'tauratiens
`
.
0. 4-1.2
REUIEMNT RQIEET
General requirements arc applicable to all drive-1 systcm configurations. There also are specific requirements that vary according to th~e aircrafl configuration and intended mission (par. 7-1.1, AMCP 70)6-201) and general rc. 1uiremcnts peculiar to particular configurations or arrangements of engines and rotOrs (par. 4-1.3). 4-1.2.1 General Re~quiremnents Ccrtain requirements are common to all Army helicopters regardless of configuration or intended usage. The desired level of attainment of th~ese requirements and their relative importance generally arc specified in the appropriate prime itern development specir-cation (PIDS). Because ne transmission and drive system represents a significant portion of the total complexity and cost of the helicopter. these common requirements must be considered during 4-3
A
P
•-detail
design. Such require ments -- without conof their relative importance - include perreliability, maintainability, and surviva-
output, with smaller amiounts of torque being carried. as the distance from the output incmueses. The wecight of any gear reduction stage is proportienal to the second or third power of the torque. Therefore. itorth-
-- • 4-1.2.1.1 Perfoima|e' The contribution of the drive system to heclicoptecr
while weight savings can be made by using drives in the order ranked -- concentric drives near thc output, parallel-axis types at intermecdiate or conmbining
performance can be defined in various ways. lHowever. the folluwing factors predominate: 1. Weight 2. Efficiency Size 4. Noise le' el.
stages, and intcrsecting-axis types nearest the engine if drive direction changes arc required. There are occasional instances where these rules may not apply: e.g., high reduction ratio may be undesirable at secondary-lcvcl power outputs, such as tail rotor or auxiliary propellers. because of the long distance from the helic:optcr CG. Althou~gh thc total drive subsystemn weight may be less with a low'cr tur-
Ssideration Sfoimancc, bility.
S3.
-0t&-
-.
4-1.2.1.1.1
Subsysitem Weight
Weight of the transmission and drive system is minimized by attention to compact size toge~her and
with the use of high-strenlgth materials for dynamic
qie being carried by shafting, it may becoine difficufqt to obtain a satisfactory CG location due to the larger nmoment of the extra weight of the higher ratio final reduction stage. The specter- weight,%of current helicopter main gearboxes in range from 0.30 to 0.50 Ih/hp for red uctio n r a tio o f 15 :1 to 7 4 :1 ( F ig. 4 -1) . T h e n ext uc •:auc U ||uuubtic uly w il . ... this lrndlcx dr "oin.. .... the 0Z2 level. 4-IZl. TnmsonEfcey(
components that make maximum use of the material properties available within the limits of pr-rmissibic failurc; rates and reliability requirements Superior . = . . oLc.~ . . .. . - -A.. a it . . . o .l ,M:. ~ ~ ~~q ..... low -d ensity m aterial for forged or cast housings also necessary. The latter is import-.nt because gear:'%•are l•_-• ~ ~~box housings comr.prise from 20 to 60% of total transmission weights in current designs.4-.112TrnmsinEfcey The requirement for efficient power transmission is large gear ratios per stage generally SExcessively of such importance that gear types other than those weight. The pinion size is determined primarily •'••.•add Slisted in pai. 4-1.2.1.1.1 seldom are considered -- ? by the torque. transmitted and is relatively ind¢-• for application to the re-Ain po%%cr drive sc~iou.;y gear of the the weight but star ratio; o1' the Spendent member increases roughlyus the square of the ratio. Large ratios per stage usually also are inefficient. mConsidering the sib of support structure and bearoor ings, as well as of the gears thmselves, the gear types themselvgs by increasing weight and power loss worder 1.(for approximately equal gtaonratios) in the follow-cri.t ingThmanner: Concentric drives: epicyclic or planetary devices falr4a4 n ciblt eurmnsSpro Star - ~uiiycotrlinpr~~sngad reversing 2. b.Parallcl-axis drives: rotation. spur, helicalo *Ii'!and horring-f
h-
ris
pt prlxd gabxsi ag rm03 =ducionrtoom51t sg n itrct-a 4:t
o05
bh
o
u
e
r H.4I.Ten
bone
3. flterslwing-axis drives: spiral bevel and hypoid.reurd This ordering reflects the addition to tor-thcs the intended speed change per effects, stageW ofin additional
/,/
"
c-rules*D
.que vector translations and rotations. The con-
S|centric
axis; vhe does nota translation alter tht torque dri-'€ .paralll drive requires of output vector dre with respect to itruto and the intprsrcting-axis driveo-ellers, introduces a new coordinate tfrough tht rotation of the output vtttor with ocspect to input. The selection of drive types should follow the given ranking, beginning at the ouput drive. Also, the plargst reductions should bw•ith kn cloto to the final
Q,
on ""--AH'IG
b
of t
TActEOFF PO*FCR lopt 10" F jure 4-1. Helicopter Main Gearbox Weightos Takeoff Power
" -
J
.:
-
i..
,
..
wc~-I 7*2 " example, with a 1% in-. "train for many reasons. For crease in power loss the life cycle cost for an assuned fleet of 1000 medium helicopters would be increased by $100,000 per helicopter by the extra fuel necessary to perform a constant mission. This is based upon a
6000-hr life, a specific fuel consumption (SFC) of 4 0.65 'b/hp-hr, and a fuel cost of $0.016/lb. Firthcr, the average helicopter lift capability ranges from 5 to 10 lb/hp depending upon rotor disk load.ng and 3I3 C. operational variables. Therefore, a medium helicopter of 2000 hp suffers a useful load reduction of fiom 100 to 200 lb with a 1%additional power loss. Thus, when comparing an alternative gear system of 1%lower efficiency. the basic gearbox weight would need to be reduced by more than 100 lb to compensate for the power loss. The reverse drive efficiency of a gearbox also must I 303 2o00 1000 be watched carefully, as excessive use of recess action gearing can create a problem in autorotation. For inENGINE INPUI POWLR.hP stance, Ref. 2 dcscribes high-ratio recess action Figure 4-2. Power Loss to Heat vs Input systems that operate at high efficiency as speed rePowIm - Typlcl Twin-eMgime-driven Gearbox ducers but become virtually self-locking when operated as speed increasers as in autorotation. A helicopterautootatonaldesent of 15,OUU tlb200 grossftminis weight making an •singits as those shown in Fig. 4-2, as possessing a given efautorotational descent at 2000 ft/mmi is using its ficien4CY apart from a specific power rating. Tise transavailable potential energy at the rate of about 900 hp. mission loacs shown result in efficiencies as. follows: Drive system windage. flat pitch tail rotor drag, and a In l ownr. Efficiency, few minimum accessory loads require appro.6imately Input ower, Efficiency.% 100 hp. A 95% reverse drive efficiency leaves about hp 7 795 hp, a reasonable level to sustain the prescribed 500 74 descent. However, sudden yaw control requirements 500 93.4 conceivably could boost the revirsc drive require1500 96.9 iment momentarily to 200 hp. In such a crse, either 2000 97.4 the rate of S,1.:__:. . .descent . ... would . ..increase . ... 4sharply r . .hor ..some ; 2500 97.8 Iw111i.t,1c crll.gy outllw 'I 'borrowed-frli.the maii.98.0 3000 rotor, with aby slight reduction of rotorvelocity. speed Howbcng Th0windage losses are influenced 98.0 strongly by oil S~~compensated an increase in descent copesaedbyaninreseindecet elrcty HThe ever, if the reverse drive efficiency were 50%, the tail rotor and accessory load would extract 400 hp from the main rotor, requiring almost 50% increase in rthe mfadesent r torrqusing an an almost50 rtr in seeds, rate of descent to sustain safe mnain rotor specls. The power losses of a typical high-speed twinengine-drive main gearb x operating at constant speed might vary as shown in Fig. 4-2. At full speed there is 25-hp windage loss with zero power transmission, and then the power loss due to friction is added as the power transmitted is increased until a total loss of 60 hp is reached for the full twin-engine power input. The slightly downward inflected curve shape is rather typical for most modern, heavily loaded, high hardness gears and antifriction bearing systems, although in som- instances a virtually straight line may be observed. Clearly, it is improper to speak of a ScarboA having loss characteristics, such
viscosity, the amount of oil supplied to the various gear meshes and bearings, and the oil scavenging characteristics of the transmission. Small gear-tohousing clearances, poor drainage paths, and excessive oiling should be avoided. Good estimations of gear windage losses F. may be obtined from Eq. 4-1 (Ref. 3)
P. -
10,, 100)( 10
hp
(4-1) )
where D - gear diamater, in. L - gear face width, in. n rotational speed, rpm This equation represents an application of basic propcller theory (Ref. 4) and is based upon air density at 4-5
"
.
'i
AMCP 706-202 standard sea level conditions which i, 0.00238 slug/ftV. However, for W.IL-L-7809 cii at normal operating temperaturci p -1.748 slug/fil. Consequently. if thi. hel~coptce designer can estimatr. or experienirtally determine the oiliness of the -. nsmission atmosphere, an average p may be employed. Ref. 4 saggests 1, )4.25:1 nir-oil ratio in which case Eq.. 4-1 can bte exprs"cd as / P,. - 2.18(
DI
,
hp
(4-2)
10"-F
Methods of gcai mcsh oiling also affect windagc lossesi 1)iffcrnces of virtually 200% hmve been reported to exist between in-mesh and out-of-mesh oiling for a large. iinglc-reductiorm gear set (Ref. 5). The torquc or load-sensitive centribution of the transmisision to power loss is duc almost entirely, to the rolling/sliding lo!nd-carrying bearing and gear cdcmcnts. Good first-order approximation, for antitriction bearing tosses it mioderate speeds arc given in Refy. 6 and with a little greater precision in Re .f. 7. However. where very accurate decterminiations arc required or where high-speed applications exist, the prediction of these losses requires an understanding of the rrr complex factors involved (Ref. 8). For optimizat ion studie~s and examination of the effects of external dcflections and asymmetricAl loading, there is no substituic for a good digital computer program of the baIc equations such as is presented in Ref. 9. In general, the significant power losses in bearings uc~tut in regions of appriciable .ýontact zone slip. uýSIC re the most efficient for most applications. Radidiy loatded ball bearings are lower in efficiency. whl b, aular %cntact (thrust loaded) ball bearings C.'N64~ sivi~iti.,i..rty greater friction losses. When the 1; latar aft: used in verry high-speed operations. their r'-iwem 11-)3characteristics dcttrioicite sharply as ex* cc~silve :ctifug.11 and g.-yroscopic forces affect ball ýUICeM-tiCl. SItandurd tapcrc.1 :iollr bearings (conical tolling ticincints) ire the 1kiht zfficicent of thesc fout. typres. ia~becaosc cf the hea.'ily loaded cone rib that'.ý in id~contact whiit ;he largt cod CS the i,ccnical roPliti. A. ti,- type os' anpuhat coniact cylirt. drical rci.er (Ref. ;01 rmi), offir hdvami.ites when fuloy dcvetoped. In Nhs ty~ve, ?ý)c co*ne , b lo~'d is theorcti,_-Oy reducA;J. bjh iincreascq s-lidiril' rttsuIs V~ nCm or Wt~h of t'ic -jic-rellcr co-aaci~s- Expcticivre Is '.nsuff!.icnt f,ýr L-valuel,,i'ol &Itz. relative iicat utnrnAtioil. bitt irdicatio, -A are O~a! it wit! pyrov' it,, b::
.'
A-.re
effincrt, than iht. staaCurd tupcretJ zolicr znd
that itA !osses will approach those of the arigular co.)tact bitll bearing. The plain journal hearintg is unsuitable for heavily loaded so liansmission uipplications (see par. 4-2.4.2), although it may have satisfactory application in lightly loaded accxs~ory uses; the power lou with this type of bearing is at least twice as great as for any of the previously mentioned types (Ref. 11). The hybrid1 boost bearirg (hydrostatic plain thru-st bearing in series combi~nation with the stationary or rotatifib ring of an angular contact ball thrust bearing) has been evaluated experimlentally (Rcf. 12), but no helicopter expecicnovv is known. The experimental work indticated a friction torque for the hybrid ocaring of roughly twice that for the simple ball thtust beaiing. The accuracy of the calculated losses for bearings is dcpendcnt upon proper installation and application dcsign, Excessive preload, due to insti-llation and/or the~rmally induced. c.:n camily result in doubling thc friction torque arid also will have detrimental effects upon fatigue We and reliability. The power los~s in gcaiing is ti~iiie an involved and cnrvr~ICiit neirte~efcwih ranking, listed earlier in tihs paragraph hold true fmn basic efficiency also. There are reawinable ranges of gear ratio for which specific types if grer drives arc bcst suited; when thi:~ upper limits are exceectce. the resultant power loss is apt to iincrease to a level where superior overall efficicm~y may be obtained by reventing to two stages of lowcr ratio. Tie suggested optimrnu ratio ranges are: 1. Conecrntric - epicyclic, 3:1 to 5:1 2. Parallel axes - straight spuir, A1to 2.5:1 -single
hclicals, 1:1 to 3;:1
I:A to 10:1 3. lnterecting axes - spiral bevel. 1:1 to 3.5.1. A nuintibr of assumptionas are included in these suggested ranges. The lower limit for the epicyclic assumes a simple u(oWirevcrsing) design. The planet gears bmcco-ne excessively small so that the sun driver ten&~ to uct like a specd-incrrasing drive *ith a long. inclficicrnt a.-c of a oproach. At these low ratios the epicyclic system becomes planet-bearing-capacity limited in iltht it is difficalt to fit sufficient~y large bearings to carry the nececssary load if the gear teeth are stressed to satisfactory levels. The upper limit for the cpicyclic represents a reasonably designed, four-planet idler systemt whose weak point is tht; tcndenzy for pitting of the sun gear. Be-caust the dasigni is sun-pinion-diameiter limited (to acc--ptabic Hertziart stre~ss levels) and die planet gears art; i'~ur tiiwý.5 the diameter of the sun, the. system is rapitdy becoi-aing inefficient from a weight stand-
do
06M'
_AMCP
point. Th: carrier structure, ring gear, and ring Sear housing (ir used) are excessivci,, heavy. If used in a final drivc stage, the low speed of the plane bearings x rotor speed) is not conducive to the formation of adequate oil film thicknesses for good performiance even though th~eve is virt.:ally unlimited space for high-capacity planet bearing designs. 'The upper limit of th siryl helicals is based upon grcatr than 15 deg in order to minimi7e the thrust
the subrtgimne or microclastohydrodynamics (Refs.. 13 and 14). Fig. 4-3 ieprescrits the salient cnarficteristics and interdependent variables influencing f tbroughout the rainge of specific film thicknesse. The ratioN)isintroduced to give physical meaning to thesw regimes in terms of the roughness, or surface finish of the activo tooth profiles according to
comonet o ýh tothloading. The extended upper limit given for herrirngbonce designs is based upon the use of high helix angles (4'AM 35 deg); thus. mayxim'im
where
I(9/4
Iadvantaf
h
h
6
olfl hcnspn -. mean ceiflerline average (CLA) surface
r istaken of the thrust component cancella-
tion that permits atlainnient of very high face contact ratios which in turn permit some reduction in the profile contact ratio with an attendant reduction in sliding velocities. The upper limit for the spiral be-vcl gear ratio reflects the use of approximately a 90-deg intersection of axes and is based upon increased losses due to excessive sliding velocities in both the art of approach and the recess. Tbc lower limit f'or spiral bevels would be applicable for overhun-gmounata
7UF.X
mEltrIaIIUII
Ar
4A.
aLA3IIa%,ULIaxcs
full-straddle-mounted.,9O-deg-axis systems cannot be ,) accomplished below an approximate ratio of 1 4:1 for 9'0-cleg axes. In all cases the lubrication is limited to low-vis-
* *loss
cosity synthetic turbine oils with ro unusual additives. The specific. efficiencies obtained in gear meshes are basically consiecred tu be represented by analogy to classic physical roncepts. The friction power of P, of sliding bodies in contact is given by Eq. 4-3. (4-3) P1 WV~f/550. hp
roughness, pin. Region III of Fig. 4-3 chn be ntglected fMr helicopter transmission components. The entire rcgion is defined by classical hydredynamics; and the properties of lubricant viscosity, xliding velocity. and load interact to build a supporting lubricant film) that complcte!y separates the load-cart-ying mechanisms. The observed friction is primarily dieptrident upon the viscosity of the supporting film. Region 11 actually extends (on a submi~croscopic scale) into Region 1,but the true importance of th region is that it represents a transtional phase that "nly has become defined with enigineering s-* niCi "P.n.e within the past decade. The pressure distributions within the loaded gear surfaces are considered basically Hertzian, but the Irilm thickness is deptrndent uposi the additional faýtcrs of elasticity of metals and the property of greatly increased lubricant viscosity under the H-ertzian conjunction pressures, The observed film thickness is known to increase with increasing entrainment or sum of rolling
where W
V,
load. lb
CO3
f - coefficient of friction, dimensionless Minimization of power loss would simply seem to re-
FIAL SIALE
manner with thi intensity of load [expressed as a
*
~~~and factors descriptive of the lubrication regimc~ and the lubricant itself. Lubricant regimes sit classified loosely (from thin
~i-lm separation to thick) as boundary,. elastol'ydrodynamic (ENID), and hydrodynamic.
FIRST
TA5(
-
-PLIN
-
*-8ISPFESNPU1--
LAN
quite attention to minimization of i andf. However, the apparent quantity f varies in a very complex
comprcssive (Hertz) stress Se], the slidinr velocity,
HTYDODYNAMIC
LENDCR
BOUNOAR
sliding vclocity, ft/see
OilRO1GEARS__ BARINS.. IQ~L
MAIN______ ROO
-
5-A '--
I
RAICSIIGV
I CRAC
AI'PA1111T~IFANCEPA IfCREASING VISCOSITY.
Q AYSAOBEATR!INC
lubricant
but the following approximation is uselul. In this of definition, the boundary regime includeE
-~syst.-m
-4-7
BEARING$
FIM5.5
IIS
C-HNS
AI
Figure 441. Lubricatlon Re~gimes
H'
I
i.
AMCP 706-2D2 velocities VT(VI. V, + V2, where V,, and V. are the velocities of the bodies in contact) of the loaded bodies, the lubricant viscosity at thc conjunction inlet, and the pressure viscosity coefficient of the lubricant, and to decrease slighily with increasin~g toad and V,. f he largest values of this region represent full separation of the loaded bodies, while the lower values permit some mieial-to-metal contact of the asperities of roughness peaks. The most widely accepted cxpres~ion in use today for EDH film thickness hE (Rcf. 15) is
HE tZIANORV)
LIISRICANb; L CONTACT
PRILSSURi DISTRIBUTIONI
V
/
..
LURh1
.*--
.-
-
.-
~~ V, - VELOCITY Of BODYTI
hE -
2.65
5
31,pn
/..0X
(4-5)
-
or HCOY 2 ELO0CITY
Figure 4-4. Elastic Body Contact Pressure Distribution and Interface Contour where tl'c three El-D dimensionless parameters are G =x'(materials)
u - 10F
(speed)
W
(load)
=
W~~
O t
and in. E'-combined modu'us of clasticity, R = mean transverse pitch radius, in. PT= mean rolling velocity (VI - V2j)/2I, in./%"c W = running load, lb/in. a =pressure viscosity coefficient, in.'/lb A, =dynamic viscosity, lb-sec/in.3 A physical senise for hE is shown in Fig. 4-4. Eq. 4-5I,~ is isothermal, in that it does not treat the effects of material heating in the con~anction, but is believed to hb rosacnngkv agpeutgt. I,*% f^ V /
-I
*
-
j
V
%,o1-,eS , f ,,t1t,ast
--
The observed friction in the EJ-W regime is prim'rivduc to viscous shear 4J the lubricant in the high-pressure field of the conjunction, Much experimental data exists to relate friction values to certain dimensionless parameters. Most take the form shown in Fig. 4-5 (Ref. 15). Such relationst iips hold for constant values of surface roughness and lay, and for specific lubricant types. The motimportant conclusion from these data is simply that friction is relative~ly low - on the order of O.u2 to 9).04 - for components operating in Region II. A mre detailed analysis that considers the thermal aspects of 131-I1) solutions as applied to simple involute gears may be found in Ref. 16. Region I, defined as boundary layer lubrication, represents conditions that predominate in the lowerspeed com,'oncnts of helicopter gearboxes. In this 4-8
3
O *
.v.k
(d. v V, V.
Figare 4-5. Friction Parameters-
Coefficieni vs EHD Region 1andi k
region, f may be influenced significantly by interaction of asperities in the rubbing load-carrying elemeonts, be they gears or bearings. The thinnest of films represemited in this region may be monolayers of lubricant products that either adsorb or adhere to the extcrior molecular s~irface of the metal. The variables influencing friction include the chemical composition and the interaction of the metal and lubricant combination and the roughness, lay, and texture of the surfaces with respect to their rubbing directions. It is at the lower speeds that very noticeable differences exist in observed friction between the arc of approach and the arc of recess of involute gearing. Fig. 4-6 depicts a very low speed mcasuirement of this pl'.rnomenon involving a spur gear set of minimum atiAinable profile contact ratio (CR) (Ref. 17). The
Although the cxpenimcnts cized were. conducted in bcwundary lubrication ccnditicns that yielded much higher f-values for teeth of coarse pitch. it is int twresting to note thiat the f-valuc obtained for gears of finter pitch was in the range of values expoctod for El-D lubrication. Thi3 iliustiatcs th-t iinportvaice of using gewring of relatively tine Vi'tch to obtain maximum efficiency. In addition, it should be not'nd that the apparent differences in friction betwoens the awvs
very low CR is employed to study the extremes of these approach ard recess portion effects without introducing the data confusion that normally would ocin the zones of double tooth pair contact. Fig. 4-6 *cur also illustrates the driematic improvement in f that results when the contract ratio is increased. The torque Q and pitch diameter D., being held constanit, the higher contact Yatia was achieved by changing the diametral pitch Pd aind the number of teeth N.
-0CR
-~
CR
a
10
Q 130 lb-in.
Torque
:~i9.11i
16
_N
-0.14
1.74
4
0.1d1 2.
1.03
2I,
_
_
_
_
p._ _
5.85f
5.55
rpm
~~1-
40
_
.
.-) CD,
~0.06 0.04_____
0.02
-12
4 0 -4 -8 ANGLE OF ENGAGEMENT , deg
8
NOTE: AVERAGE f 7 0.081,
12
10.034
Figure 4-.Angle of Engagement 4-9
F
*...i II
of approach and recess art masked completely by the avraing effect found in the zones of double tooth pair cont&ct. The basic tiends of friction change in approach and rcs action arc still valid i:i lubrication Region II. Fig. 4-7 (Ref. 18) represents experimental data taken at corasiderably highti values of ourface sliding velocity, load, and lubricant viscosity. Alth..ugh the use of a dig'tal compute: program to examine the many instantaneous contact conditions that occur b. a pair of gear toeth rolls through mesh is the more precise meitod of catulating efficiec:cy and studyiog detail design variations, c fexrek-i -t - -may be obtained by using average values for f. The percent power loss P, of a gear mesh is expressed (Rcf. 19) as: p,
L X 100 96, Mfeatures f
(4-6)
where M = mechanical advantage, dimensionless Values of M for various combinations of pinion and gear tooth members are listed in Ref. 19. For the ,4) o Fi . q-6, ahe •i ia: lar ( 4 u, ia~h ' , . 4.6
P
X 100
1.76%
(4-7)
II~
and for the fine pitch gear set (P4
10) the P, is:
2-.034_ Y0100 - 0.40%
/' P,
(4-8)
8.5 Their corresponding efficiencies V are then simply ij 100;- P, and we find. 9 = 98.24% (4-9) - 99.60% where the subscripts c and f indicate coarse and fine pitch, rexpectively. The frictional differencer noted for approach and raxwzone of involute action are characterized, with respeci to the driving member, by the rolling and sliding contact motions being in opposite directions to one another in approach but ia the same direction during ro.ms motion. The scensitivity of friction to the lay and textural of the mating members in lubrication Region I is s.jown clearly in Fig. 4-8 (Ref. 20). These data represent the results of experiments conducted on a gearedisk test machine with 3.0 in. diameter, 14.0 in. crown radiua,, case carburized and ground, con0 93a c . s'-smab!e e•iec ode vac . n steel disks. The circular ground data were taken with disks ha%ag a circumferential finish of 8 pin. and an axial finish of 16 gin., while the cina-ground ditks
[0..
4O~
~
-ROSS
" •
GKOUN9
300D --
0.%
3.'
''%--
0
500
lDO• 1500 SLIDING VELOCITY V..
2000
ZSO
±
Emm
--
--
SLIOINCVELrClI Y V,, -M.k,.ۥ
Figure 44. Effed of Surface Texture and Figui 4-10
4-7. Coefficent of Frkilton vs Sliding Velocity
Lay on Friction nd ScffSag Beh• vr
"
V.
*limit
-~
bad a drciinforweti ftaub of 16 pin. and va axial finis of Spin. Conm'queud.ty both types had identicaO reduced finWis number &-values and benice vii tualuly identical A-valies, but the cross-ground data news was RcEO463 irn both type., and the cross-pound disks weas prepared using grinding techniiques nor-naqly used for spur gear Waoth manufacture. A constant ratio of IV/VT - 0.556 wias represented, sai the lubrication was jet auppNWe MIL-L-7806 at 190*F. Theretome a particular point on agear mesh wher V, 3S.6% of Tr is represntmd on this ftzgmat u a linearly increaed gea apecd (rpm) by moving from left to rig1bt -ýith inremasing Y1,. Unfortunately, then. data do no reflact a oonstant load, but rathar the baa6 as delrined by scuffims. The designer shuld be aware that udrfxd lubricant condition there is little he can do to control Sat low speeds asid from refining the surface finish, Additiona-i powe: lowa sources of a transmission systemn include the accessories anid the oit punip. Accesory power requirements are fixd by the individual helicopter requirements. the exact typc of accmoasy 5nvolv')d. and the demand or duty cycle memiPaAMtmnfib
mmrv u
,nwewr
vAmnnrap.
ments igdiscussed in Chapters 1 and 9. Oil pump loss P,, is estimtted adoquately by the aimple equation: Fpk '7 (41) PPhp qP where oil flow rate, gpm discharge pressure at p;Anip outlet, psig converion constant for units, 5.83 X 10-4 efficiency of pnmp (generally from 0.5 to 0.9), dimensionless For a 20.gpm system with -jregulated discharge of 60 psi, the pump outlet pressure would be 120 paig under typical cofiditions. For an assumed pump efficiency of 0.5, the loss would be: (20)(120 X 0-')result (5.3 (2)(10.) (59 1-) 2.78 hip (4-11) (0.5)through F
-
p kr q. -
4.1.2.1.1.3 Size Compact gearbox size is important in the achievement of low subsyst;= weight bemause Ibc housing or casinj that enclome the dynamic components contributes a significant proporticni of the total systam However, compaction should not be emphasized to the point of causing excessive oil churn and windage loome to the detriment of efficiency. Ref. 21
)weight.
suggests that cide clearances of 0.5 in. or pester between Seats and cauing wall* reult in negligible loom due to oil churn. The required clearance[ between casing wail and Sear outside diametpr inarc of conformity, speed of &ear rotation, oil viacosity, amount of oil etted on the gear mesh. &nW the amount of run-off or drainage ouil in the locattion at question. There are no formulas for calculating satis-A far~ory diametral clearances, but some successful design applications have amployed valu&t of 4-. proximately 0.5 in. for 2000-fpm pitch line velocities and 3.0 in. for 25,0I00fpm vcocitics for 1801 dog of conformity and kinematic viscosities below 10 centistokes. Evtn at these clearances. it firequcntly becomes necessary to provide scrapers or some mean to retard vortex acneration and lo Waized recirculation of the oil. When wet sump systemns are employed, suffic.ient vei-tical sprce must bý provided to keep the operating oil level (iaucluding th4& Avr~iawd or foam lazyei) Weow the gears and beerins
i
j
4.1.2.1.1.4 Noase LereleV The fourth jxcrforrnince criteria of low nao" iceW &--&a2--
~
--
~---
-
noise is 5Isuailj of imporAncc cidy ii, relation to crew, and passenger comfort levelt, whale rotor ttid enoine noise arm the principal contribute,-& to thc auiral dotectability of the helicopter. The funda.mental gearmeshing frequencies. which range from 40 to 22,000 Hz in present-day gearboxes, are the p'impty sources of noise. Refs. 22 and 23 identify thc magnitude of the problem for two helicopters. Tht overall helicopter configuration and the rfaulting number and location of gearboxes dictate the areas nlTfeed by noise; e.g.. a tandem-rowo: helicopter may have its forward transmission located above the crew compartment. resulting in less a'avorable noise conditions in the passenger area, while the reverse may be truer gabxa for a single-rotor configiuration. Gear noise may emanate from the gabxa of forced or resonant vibration of the housing or cases. It then reachca the.-iz-r passengers either direct airborne paths (window&, access panelk, or door seul) or through airframe structural pathways connected to the gearbox mounting system. It is far more efficient and desirable to combat suo& noise at it&source rather than to rely sdely upon the use of insulating and soundproofing coatings or blankets in the crew or passenger compartrients. The latter measures generu~ly add more weifjht than would be needed to make comparable improvement in the problem at its source. Sound insulation also in-I cref sw4 maintenancet man-hours duc to the need fot 44.1
*
1
temoval Of the material during airrrame liaspections.
Also, soundprooflng efforts often are deleatod when torn or ciI-soked mateirits art removed and never ri~eplcod, The uae of elastomeric isolation mounting devices at the gearbox and hanger bearing supports is highly effective in rnduc'ang structural noise. Airborne noise should be minimized by clintisating any housinp or case resonance through use of proper wall thickness, shape, or internal gear or bearing quill attachmcint methods. While it is virtually impossibl* to calculate these conditions with suffkienrt accuracy in the design stage, they are itlatively easly measured during initial component testing, and corrective redesign then can be undertaken. Modifications in the shape or involute pr ofiles so zs to change drastically the fundamental and harmonic noise content have been investigated analytizally (Ref. 23). However, the slight variation in profile requite to achieve theoretical improvements was judged boyond tLn. p~resent rltinufacturing state of the art Soni't methods for approximate analytical prediction of resonant per* formance for relatively simple structural housing * :ha~ L~ dvaniMd an RC.24 It is not certain that the elimination of all interactin& vibratory and resor..rit behavior in the various gear meshes Is entirely beneficia; i.c., some sacrifice in the efficiency of the lower speed (boundaty lubrication regime) meshes may result. The effects of axial lubrication upon the reduction of tooth-meshing friction is reported int Ref. 25. Although the engineering field of gearbox noise generation is imprecise as yet, there exists considerable general knowledge that can be of practical benefit to the designer. For example, it is known that hnigh contact ratio gearing and finer pitch sizes pro. duce. ess noise than their opposite counterparts. Similarly, helical gears ame quieter then ttraight spurs; spiral bevel gear are quieter than ctraight bevel or Zerol gpars because of their greater inherent contact raitios reduced dynamic increments or waste loads, and increased smoothness of operation. In addition, increased gear tooth backlash and clearance can help to maintain subsonic air ejection velocities fromt blgh-sprtd meshing teeth (Ref. 26). Viscous films between tAtionary bearing rings and housingsi can provide suffikient damping to roduce vibratlon and noise propagation. Coulomb or dry fric* tion devices have been successilul in damping resonat modes in Sea rims and webs, as have high hystaresis materialis clad or bonded to shafts and webs. onthe 4-12.1.2 UdhIEfy A complete general discussion of reliability con4-42
cepts is contained in Chapter 12, AMCP 706-201. Thin paragraph, therefore, deals with specifics as related to mechanical power trantsmission ccmponents of the transmission system. Concept definitions and numerical values for quantitative reliability indices generally are speciriod in procurement documents and, with increasing frm quency, in PIDS. For transmission and drive system design there are two types of indices: I. Values for such characteristics as mission reliability, flight safety reliability, and system reliability rot the entire helicopter (us.ually for a gi-een mission and operational enviroanmcnt). Typical values and methods of expression might be, respectively, 0.90 to 0.99 for ont. hour of misL,.ion time. one failure per 20.0(N) flight hours, and 0.70 to 0,80 probability per mision hour of no system failures requiring unscheduled nainteniance. Because the reliability of the helicopter is a composite of the reliabilities of the individual subsystems, individual reliability levels must be assigned %stargets for the design effort. As an example, assumei the Request for Picoposal (RFO) specified values of 6.98 and 0.9999 (one failure per 10,000 hr) ur minkimiu and safivt
criteria, respectively, The Ai-
lowable apportionment for the drive system would be dependent upon the complexity and type of helicopter, but typical values for this apportionment could well be 0.999 and 0.9999. respectively. Techniques for desigri to these requirements will be addressed in par. 4-2. 2. values for suhsy,6tems and component~s for such characteristics as reliability after storage and mean time between removals (MTBR). Typical values are a maximum of 10% degradation in mean time betweenA failures (MTBF) after storage for six months in approved environment or containers and 1IS0 hr MTBR, including both scheduled and nonscheduled removals. This type .-f index is directly applicable to individual subsystems, including the transmission and drive system. The MTBF values subject to the 10% maximum degradation limit lite those specified implicitly or explicitly in Item 1, i.e.. the 0.90 to 0.99 mission reliability carries the recirrocal meaning of a 10 to' 100 hr MTBF, the one failure in 20,000 hr for flight safety is a statement of 20,000 hr MTORE, and the 0. 70 to 0.80 probability of zero system fail urs per hr implies # 3.3 to 5.0 hr MTBF. The MTBF levels corresponding to the drive subsytcm apportionment in Item I are 1000 hr (0.999) and 10,000 hir (0.9999). respectively. It is important that the designer understand that MTBR arid MTBF criteria discussed previously ~ interrelate in a unique fashion with the subsystem design reliability when a finite timne between over-
(
'
(
.
F)
A&IcP 705-=2
hauls (TBO) is selected. On the astiumption that. n~l transmission and drive system failures are of sufficicnt magnitude (and detectability) tn. force a mis-
cancellation effecta, leaving a not MTSIF value dofined with sufficient precision to express design reliability rc-quiremenrts.
to be aborted, and further that no TOO (scheduled removal) reqjuirement is imposed, then it follows that the MTBF owMTBR. The imposed requirement of a 1500hr MTBR would necessitate a failure rate A!• 0.0O004 or a one-hr mission reliability of 0.9933 for the tra 1smission and drive system. The relationships satisfied in the preceding statemenits are:
Yh "ft of a given set of gearbox failure data to teepnniiasmto eel eti hrc tcrestxcsofth asysumtemion reveals. cfteoertaingcar loprtng, a If sufihenl syroum in*q teimeci ofah cetwerotrndwlb bevdasaodfnt erottedwl eosre sann cnsatorireigfiuert.Omoroetisensilive (wearing) compontnts eventually will begin to dominate the failure picture as the end of their tueful life is approached. Generally, those componenits operating in deep boundary layer lubrication regimes W19I be the first to influence the picture. lit such instances the wear will progress to a state wherein conditions become fa'vorable for the occurrence of a failtire mode indigenous to that new set of operating conditions established by the partict'li wear state. Fig. 4-9 gives an example of a well-developed or debugged gearbox with relatively low TRO that satisfies the rancom failure characteristics a:curately dcflned by the exponential distributi.on (Rd. 27). Fig. 4-9 represents a gearbox with a 350-hr
Ision
I h and using Maclaurin's form of Taylor's formula R xp~(k MT
.B
(4-12)
where R =reliability (for a I-hr mission), dirnensionless However, if a 2500-hr TBO level were to be assigned, the I 500-hr MTOR could be satisfied only by a higher reliability number (lower failure ratte). The rclation. . - - --
~bnp
Imuy uti
-A
flf IL-. L.. U..f
rpirbbvd ivi bi.0-ha, ivjjaffuw.
3MB
_A< AAIBR
-
XTaO
-
-
0.00067
___00
--
o.00040 X ý.o
•! 0.00067
-0.00040
) -
sn..
ii
. Cý
fý
a50
I
1,0l
-
I
-1
_
scheduled removal frequency, the resulting MTDR is (4-13)
2500
Therefore, A. !5 E •B~ :5 0.0027
scheduled tria
reliability of 0.9930. However, due to the 350-hr
TO+Xr1 B
ABO+IM
AMBR-£ =
_..--
as.
lic~cthcnc-v TBFmus be~ 304hr riditecor IT# I "A responding ireliabilty R ;± 0.99973. It is paradoxical that such factois as flight safety reliability and the increased cost of overhaul of a badly degraded gearbox (extensive secondary failures) may fix the TBO interval at a level that in turn rm quires a significant increase in the required MTBF to achieve a specified MTBPR. Relationships defined in this maniier are tacitly assumed to fit a simple exponential failure distribution. This not only simplifies the arithmetic involved, but enables the designer to think directly in terms of the inverse relationships between the number cf detail components comprising the sub-. system and their intrinsic failure rate requirements. Although many individual gefirbox componients are better represented by other distributions - e.g., Weibull, gamma, and lognormal - the averaging effect on the subsystem as a whole is such as to invite
204 hr.
Fig. 4.10 from~ Ref. 28 represents the distribution evidenced by a %i~el dc-bugged gearbox with relatively high TBO. In thin caa- the upward inflectiorn or concavity of the data points reveas the strong influence of component wear-out as the TBO level is approached. ISimilarly, Fig, 4-10 represents a gearbox with an I100-hr scheduled TBO, an MTBF of SUP' hr, and a one-hr reliability of 0.9998. Ir~ this cast the scheduled removal frequency results in an MTBR of 904 hr. Obviously, design MTBR and MTBF values sart dependent upon the desiner's knowledge of repro 5entative failure modes and his ability to assign reasonably accurate failure rates. Extensive test and service experience with analogous componewts and subsystem elements is essential in arriving at realistic predictions. There is little published literature to aid the designer, but the selection of components with known lower generic failure rates always should be the objective. It should be recognized that generic failure rates (indicativz of the intrinsic reliability characteristic of any component) cannot truly exist apart from the envirvnment in which the component functions. Some insight into these environmental inlucrnces is given in par. 4-2.1.1. 4-13
&mill
-
0
--
GEAR BOXES SURVEYED -.173 32
-
32
26TB -
OF FAILURES . 34
26 TOTAL TIME
-
-A
-HNUMBER
-
172,431 hr
,f71ti--
10
0-0---1
00
01 0 00
-ii
0 12________
12E 17.67
4T TOA
0
OPERATIONAL EXPOSURE TIME;, hr x 10-1
Number of Failues a Hewns Simv Overh. - MATSF ow5001 r
FjS449.
2r
0
0W8
-'7-
04 121010 2
44
6
OPERATIONAL EXPOSURE TIME, hr x 10f Flgurs 4-10. Number of Failures vs Hours Slae Operatin. WINF - 9W br
based upon th. profile of !he design mission; i.e., for Satisactory failure tate estimates for rolling elea mission profil, of n discrete values of bearing load nment boarings may be derived by seeral wethods. P, each occurring for a perorntue a, of the total one simple but sulficient value is: (41) operational time: IN-
S L
I
5 /~h-I(414
1/3
probability of siurvivl, dimensionless AMC i(4.15) - duuign life or scbeuled renowvatime 100)l (TDo) br where The value for Sis rsd fromthe curve in Fig. 4-11~U corrseponding to doe rtio of dbo design life L to thw0 Although in ma-iy instances the life-load exponent life olo at wbicb i0% of the beauing population will more correctly can betaken as 4. thecube value is rad Tkin#10 value for a gienbern gis determined recommended for determinations of failure generally root for the from the bearing msnufacturW&' data rate A. mean cube AMC load. The applicable AMC load is 4-14
-
(.
XA&P70&202
-
0.6
will have failed by emrplo,41ý'a dispersion exponent, tatii'c of typical h fcopicr geat performance whome
I
4 Ii *1
excellent quality control generally results in lower 6ispersion. Very steep slopes frequently Lubrication Regime 1. The life value L2 for 98% reliabilty may be read from Fig. 4-13. a Weibull ~plot. The resultanit failure rate A is then:
*population
0.4
1
1typify I
(4-16)
X - 0.02/L 2
The mean value may be taken as the 50% or median rank foi such steep slopes without losb of significaid accuracy in using Fig. 4-13.
0.1 9S.9)
99.9
go 98 14 9 9..4 99 PROBABILITY OFSURVIVAL S.
9)
94
90
FSpot 4-11. Pro-babillly of Survital vs L/810 Ratio
-
2D
copter appfications will exhibit pitting as the lifelimiting failure mode (pat. 4-2.2.1). The life-stresa-to-4. lationship for gear teeth is far more complex than for arc ar ighe, more types of ni~ctals and heat treatmctfare prgevaen, and the elastohydrodynamic and cemical effects of the lubricant are known with less precision Available simss-hfif curves more often than not are aacd upon the mean pitting, or spalling, endurance of an unknown statistical sample (Ref. 29). Intensive is underway by many organi7.ations (ASME R esearch Program on the Relationship of Lubric8-
I
____
IS
01I0 PLIGiecls
Figre 4-12. Spalling Life vs Hertz Stress WEIBULL SLOP", 5.5 -1
Kresearch
-
-O
ton and Fatigue in Concentrated Contact, for
70
preparation of exml)that should result in thetrdifi"ehr
60
N'~~~ pit-1i
that consider lubrication regimes, materials and metallurgy, and sliding speeds. The AGMA data of Ref. 29 reflect usc of a stress-lifec exponent of between 9and 10, whereas values of 5 have been reported (Ref. 30) for operation in Lubrication Regime I (Fig. 4-3). However, in the A!scnce of well-documented, statistically-significant, test date, the AGMA data should be teken as represcntativc o~f most gear applications. Use, of the RMC value of the Hertzian or compressive stress in the contac! area to obtain the
1tos-1
4o. ---
~
*j
4-
1
'i
-
e
23 I
10--f.-I
4
-
mean spalling life from Fig. 4-12 is satisfactory, al-r though the quartic mean level has be'n shown to offer excellent correlationi in Lubrication Regime 1.
7)
There are many suitable techniques for reducing
this life to a value of failure rate A.One rather simple method uses standard Weibull paoper to reduce the mean life to the level L 2 at which 2% of population
-
Th0------ T
200
Geat failure rates can be determined similarly from design stress levels. Properly designed gears in heli-
gK
2GEAR! GRADE 411.02 AGMA
300-
-
2-
10'
0 $PALLING LIFE. cyoles
Flpre 4-13. WelIulI Me - Spalifq Life vs Gear Popula"o Rank
0
AWP 7Q2
-
41-1,1.1.3 Maaimalahlty A gcnoral discussion of maintainability may be found in Chapter i, ,k MCP 706-201. This discussion, therefore, treats considerations relating specifically to the transmission and drive system. The basic concept of maintainability often is expressed as a requirement for a specific or maximum number of maintenance man-hours per flight hour (MMH/FH). Army helicopters of a few decades ago exhib~ted values as high as 35 MMH/FH while helicopters in the present Army inventory have values ranging from 0.5 for the OH-58 (Ref. 31) to 6.5 MMH/FH for the CH-54 (Ref. 32). Small helicopters as a rule show better maintainability values than the larger, more complex machines. However, the values for any given size of helicopter may vary by 300% depending upon design variables. Current RFP requirements arc in the range of 5 MMH/FH for medium-sized, twin-engine helicopter for organizational, direct support (DS), and general support (GS) maintenance levels combined. A value so stated must be apportioned in turn (par. 4-1.2.1.2) to the various subsystems to establish individual design goals. Achicvcmunt of sadisfactoui-y -aitai•-ability levels is dependent upon two factors: I. High component reliability (par. 4-1.2.1.2) 2. Ease of maintenance, Mainteinance is generally thought of as comprising two categories: nonscheduled (due to random failure or accident), and scheduled (due to time change of
or clamps. !t also should be emphasized that true leveling of •lte helicopter is seldom achieved for comnponent change at the direct support level. Therefore, when heavy components neceLitate the use of hoisting devices, extra care must be taken in the design of algnment devices and structural cicarancs, so as to reduce maintenance effort. External shaft seals always should be essembled in easily vemovable housings or holders to permit bench changing of the seal element. Squareness, alignment, and cleanliness practices all are critical to the proper performance of a seal and are difficult to achieve when the seal element must be changed in place. The shaft upon which the seal operates also must be easily removabie because good practice requires that the shaft be slipped into the previously installed seal, allowing the use of adequate shaft lead chamfers to minimize the danger of seal damage. It also is desirable to have the shaft engaged with the driving spline or some other guiding device prior to making contact with the seal to prevent excessiv sidc loading of the seal. The attachment of all external components should be such that one man can remove all fasteners and similar items. Two examples of poor design that require unnecessary manpower are: I. Bolt-and-nut fasteners through structure where one man cannot reach wrenches on both elements. Tapped holes on nut plates are proper solutions even though a larger number of cap screws may be required because their rigidity or strength may be lower
wear-out components, interim servicing, and in-
than that of a bolt-out joint.
spections). However, all maintenance. concerned with componer.t change is discussed herein as a group. The generic failure rates of many external com-
2. Components that one man must hold while another installs fasteners. Possible solutions include the use of guide pins, longer pilot flanges, slotted
at,,ch Donents in the drive sytem rre
,-.aran-.,, hnl,.
that !h, cnm-
r-t,,nn
n. ,rt
-r.n..
.. h
". •-
r.;•
portents may be ranked in order of required fre-
tion devices to secure the component temporarily.1
qucncy of removal or adjustment. Components such as hydraulic and electrical accessories, rotor brakes, shaft seals, external hanger bearings, and drive shaft couplings require relatively frequent inspection or maintenance and must be designed for ease of removal and installation. Accessibility is the key criterion Such components must not be located too close to one another, and adequate wrench clear"ances must be provided for standard tools, Subsystem components such as g-arboxes are generally not maintained at the field or direct suppori level and, therefore, they must have simple and accessible attachments with structural clearances adequate to permit easy removal and replacement. The usc of integral guide pins or tapered dowels is rccoinmended in any case where heavy components must be aligned for the installation of mounting screws, bolts,
Components that require a specific orientation to function correctly should be designed so that they can be installed only in that position, if possible. When this is not practicable, as in dhe case of some Government-furnished electrical accessories, decals may be ustd at the pad location to provide insiallation instructions. Examples of one-way components are seal housings with drain fittings, hydiaulic pumps that require line connection fitting orientation, and bearing hangers. Components that require tight-fitting pilot bores or similar devices should be provided with jacking pads for removal. One man can operate two or three jack screws (tightening each one a little at a time), whereas their omission might necessitate the use of two men to pry simultaneously on both sides of a component (and possibly a third man to catch the
4-16
r
_-A
*AMC, - -component *.
706-2012
when it breaks f(we). Proper performance of scheduled maintenance tasks such as inspection and servicing is dependent to some extent upon acsuibility and convenience, Inspections that are convenient and of a go-no-go nature are likely to be performed cin time aid with accuracy; those thut require considerable quantitative judgment may bc missed or interpreted incor-
reedly, For example, the presence of vital fluids in all gtarboxes, transmissions, and other reservoirs should be discernible from ground level without opening of complex cowlings. Min-max oil levels should be used to eliminate the nccd for topping off, and the minimum level should be exactly one or two quarts below the maximum whenever possible to discourage the practice of saving half a quart of oil in an open can. The minimum level should be such as to allow completion of several additional hours of operation at the maximum likely oil consumption rate so as to eliminate the need for adding oil when the level is near minimum, While accurate values of maintenance times for transmission and drive system components are not *
5
S
,v, ,au S•.,,t'.t ,
U.flI,,=FE,.
n,,ll) ,l.,,Wpf..,,
p-ua.aavu
data are helpful in identifying present troubleome areas. Table 4-1 presents maintenance workload factots relative to drive subsystems (Ref. 33). The TABLE 4-1. US ARMY HELICOPTERS TRANSMISSION AND DRIVE SYSTEM ONLY MAINTENANCE WORKLOAD (Ref. 23) -
j
*OR K LOAD RATING
PRO6LEMTITLE U A
VERY I II
HG
TAIL ROTOR DRIVE SHAFT INPUT DORIVESHAFT OHTAIL S ROTOR DRIVE ,SHAFT TAIL GEAROTO
ocoincident IU
I
I 4
,V
LO*4
LOW[
SY NCHRONIZ ING OR IVE
SHAF I
x
nial
with ADS-I l. This plan will include many elements peculiar to the transmissici and drive system. A good program plan requires the active participation of the
x
responsible transmission and drive system design ac-
x X
ROTOR BRAKE SLJPPORT
ASSEMBL V ROTOR DRAKE DISK
ROTOR BRAKE PUCKS
tivity to assure practicable approaches with miniriun penalties in &!ive system performance, weight, and cost.
OIL COOL ER ASSEMBLYV
-
with the reduction of vulnerability. Nor-
practice is to design a complete helicopter sur-
vivability-vulnerability program plan in accordance X
TRANSDUCERS CH 54
"INPUT
techniques applicable to this goal are discussed in
pars. 4-2.1, 4-2.4. and 44.3. Survivability in cases of combat hits is considered
OIL PRESSURE
LC MAIN GEARBOX CARBON SEALS ROTOR DRAKE SEAL AL;Y
4-1.1.4 Survlivabllly Survivability in transmission and drive syste. operation may be defined as the capability to susstain damage without forced landing or mission abort and to continue safe operation for a specified period of time, usually sufficient to rctunn to home base or, as a minimum, to friendly territory. The damage may occur from either internal component failure due to wear, fatigue, or use of a deficient or inferior component; or a hit by hostile forces. The current Army requirements generally define the period of time for safe operation after damage as a minimum of 30 min i at conditions within the maximum power and load envelope, except in the case of total loss of the lubrication subsystem; the acceptable maximum power level for safe operation upon loss of lubrication is generally reduced to that required for sustained flight at the maximum range speed at sea level standard condition. Survivability tollowig internal com.pOont failure can be enhanced through such detail design practices as identification of primary failure modes, and using configurations and arrangements to asure kindly failure modes and to limit f-ilure progression rates. Attention also must be given to the elimination or metardation of secondary failures causad by primary failure debris, and to providing for j.3sitive failure detsection long before a critical condition is reached. Safe operation with this type of damage normally can be achieved for durations of 30 to 100 hr. Design
X
I
INPUT DRIVE SHAFT Ail IG I,'LUT OUILL OIL SEAL ("1 4?I
rankings given reWate primarily to other maintenance factors on the specific helicopter listed, and are not to be interpreted as relating the workload on one bali-. copter model to that of another.
Reduction of helicopter detectability and the defeat of specified ballistic threats are important elements in vulnerability redt tion. With regard to de-
X X
X x
tectability, the primary area of concern in the cas of
the drive system is noise Wpar. 4-1.2.1.1.4). While the mower irquency noise levelsare basically associatd1
t
', ,'. -- t
*
A -
K ¾
'with the rotor and/or tail rotor and propeller. the highiev frequencies are generally attributable to the transmission and drive, and propulsion systems and their accessories, Typical Army requirtments specify a maximum scund pressure level for helicopter hover and fly-by at &specific distance from the flight path. Desge Soals for appropriate frequtrncies and sound prossuore@ are given in Table 4-2. Noise level survey iriquirements are described in Chapter 8, A MCP 706203. Design techniques to secure external as well as internal gearbox noise reduction are discussed in par. 4-1.2.1.1.4. defeat of ballistic threats must be accompliahed for smaller caliber ordnance and damage minimized as much as possible for the larger calibers. Depending u~pon requirements peculiar to the mission, the drive system components must he cap&6k of withstanding a single ball or armor-piercing .,762-mm ballot at 2550 fps, aligned or fully tumbled, striking at any obliquity at any point in the system. *,TM 75-dog solid angk of the upper hemisphere (with -Complete
larger ordnanca for %.hich damage minimization
the collector par in the main gearboc or combin~ing gearbox t-sually arc excluded from the survivability requirement by nature of their functional duplication, provided that: I.- No single projectile can kill all duplicated power paths 2. A single power-path kill cannot cause secon-t dary failure of the duplicated power paths due to firagmentation of the first. These two criteria can be tatisfied by: 1. Physical separation of the drive paths sufficient to reduce the impingement angle within which a single proj~iccr can produce a multiple kill 2. Sufficic. !size and strength of the killed-path component to attenuaste the projeenile velocity below the kill thre~hold for the second path component 3. Use of structure between the paths to confine a fragmented or loose component to its immcdiate locale 4. Use of armor to confine fragments or prevent projectile impact. eguaeu
should be considered is 23-mm high explobive: inisbMaCh otimotn Ilt n f cen~iry(HI) sighet iechnique for reducing vulnerabilty. Manly slihtconfiguratior, changes can increese surviva. he esgne to tchnqueo he avilbleto Thaspeifi Thespeifi avilale tehnque esinerto bility greatly with jut serious compromise of efw~ee the stated requirements include: fcecwiho ot I. RdundncyFor example, case hardened gears with tough, 2. Desigi configuration fracture-resistant core structure have surjplisingly 3. Self-sealing oil sump materials 4. ubrcatin Eergncy cnsidratonsgood toleranct to ballistic damage. Spiral bevtl gears 4. Armor.ec lurctocnieain and planetary gears, used effectively throughout the drive train of rimall and medium helicopters Ure in-
Y .
vulnerable to the 7.62-mm threat. Planetary ring
Aa 0
Redundancy is typified by multiple engine configurations. In thesm configurations all individual drive subsystemn components between the engines and TABLE 4-2. EXTERNAL NOISE LEVEL #X rERNAL NOISE LEVEL
IFREOJENCY. Hz BANDi CUJTEW
44.7-89.2
0.-78 M&f 709 1.410
OVEALLsosive 63 12
oc
1.4110-2.9m
(1000 .. 2000
2AZI5.833
4000)
L~~i 5,2111,222
oo
'NWA
SOUNDOPRESSURE LEVEL, CB PER CEI VE0
85 5 88 86portioned 851
76 72
CETE FREQENCY (DETERMINED EMPIRICALLY)
'ONvIFSSEM psWP
SNCILS
gears may be penetrated so that the planet idler gearsX. cannot mesh at a particular segment, but the remaining Sears pick up the overload necessary to continuc normal power transmission. The melatively high contact re jos and coarser pitch of spiral bevel gears t a facto~s that make them particularly resistant to failure from loss of a single tooth segment. Narrowface spu.- gears (less than 0.5 in.) c;an be a problem, and, therefore, it is desiroblc to use greater face widths in all primary po-wer paths. Experience. with 12 7-ns. 'n ammunition is less exten-
than with the 7.62-mini projetiles, but the same
general observations hold true with a slightly larger scale ol' reference. Glear rimns, webs, and integral
outer-race sections of planetary idlers should be pro-
such that a ricochet entering the mesh will deform, fracture, or crack the gcar teeth rather than A the tooth supporting structure. Extensive observation of main rotor gooarboxes damaged by 7.62-mm /
ball -and arurnr-pimrinS (AP) anmulinition
tv
o
AMMP 7W~202 shown the digsctive capabilities of conventional hel;copter gec~ring to be quite adequate to discharge the "spm bullet insto the oil sump without functional failure or the power transmission systrmn. Integral gear shafts quill shafts, and external intercon~nect or tail rotor drive shafts must be of au1Tficient diameter to withstand edge hits from fully tumbled bullets without failure. In thin-wall alumiinum shafts operating with a minimum of 20% margin on first whirling critical speed, an extc.-nal diameter of 3.0 in. is sufficient to defeat the 7.62-mm threat whil- a 4.0 in. diarantecr is necessary to defeat the 12.7-mm threat. Steel shafti-i may be considerably smaller depending upon the wall thicknms employed. Of course, in event of damage the remainin& portion of th~e shaft must have sufficient strength to transmit the required maximum torquc.If the column buckling torque is 300% or more above this torque, &simple shear-stress calculation of the remaining post-impuct area is sufficient. However, when the buck inS margin islevas, it isgenerally necessary to conduct real or simulated ballistic tests to demonstrate the adequacy of the design. Note that it
is undesirable to increase diameters excessively sincc vulnerability is then increased for fuzed round threats. Most ball and roller bearings are fabricated from through-hardened steels and. consoqoently, usually will fracture through the outer ring when struck by a bullet at nt.ar-ero obliquity. The outnouncling case and liner s~tucture serve to expend sorie of tha kirsetic energy of the bulict; however. cony ontional thickntsse of these structures gsterally are insufficient to prevent fracture of the bearing ring. Orientation of the rolling elements of the bearing at the instant of impact has much to do with the ring fracture mode. A zero obliquity hit between rolling elements fraquently will discharge a double-fractured "pit-slice" ring segment into the bearing, while an nligned hit often produms a single outer rikig fracturc and frequently fractures the roll~ni element so well. When the des.Sn allows, a space between the outer gearbox wall and portion of the housing supportiag the bearing liner ind ring is effective in reducing bearinb damage. P~as space provides a place for the spa?'ling debris from the initial impact to txpand and eject,
DUPLEX BALL BEARINGS
ROLLER BEARINGS
Figur 4-14. Typical Tall Rotor Gearbox
-
Vulnerable 4-19
*uRAP 706-202__
_____
thereby reducing the impact shock on the bearing Ving, Some insight into useful design techniques may be gained front examining three configurations of a sitnpie, spiral bevel gear. 90-deg tail rotor gearbox as used on most helicopters with single main rotors. Fig. 4.14 is a schematic representation of a typical minimum-weight design featuring overhung pinion and pear mountings, designated as Configuration I The pinion and gear are both in overhung mountings supported by duplex bail and cylindrical rollcr bearings. A single 7.62-mm hit on any ont of the four bearings probably would not result in instant functional failure; the boaring would continue to orcrate for some time because the considerable driving torque would bre~k up and eject the relatively frangible rjis.~ng elements and cage of the damaged bearing, Howtver, direct hits on either the pinion cylindrical roller bearing or the duplex ball bearing supporting the geau would soon result in excessive loss Of geAr
mesh position. As the ci fectivc: radial clearance of either of these bcatings increased with the resulting rapid bearing deterioration, the operating backlash of' the gear tecth similarly would increasc while the depth of tooth engagemcitt would decrease correspondingly. The probability of the gear teeth skipping mesht or breaking off upon application of sip nificant yaw control would be gkcat. Hits on thc outboard bearings would yield far lower probability of gear miesh failatrc. Configuration 2 is shown schematically in Fig. 415. The beval gear set it; identical to that of Conifiguration 1. Pairs cof the same types of bearings used in Configuration I are now used to straddle mount both pinion and geAr membtrs. In this configuration the increasing radial clearanctc in any bearing sustaining a hit wifl result in less dectrioration of the gear mesh. with a corresponding decerase in tlte probability of gear fetilurc upon sudden yaw control input. A gear-
box of this configuration designed for tail rotor
CONFIGURATION 2
~ROLLLO BEARINC-
7'
DUPLEX BAL.L BEARINGS
Flgvrt 4-15. Tail lRotor Gearbox 4-20
At
-
7.62 mm Proof
r7
AMCu±706202 steady hover power of morm than 150 hp wo-ald bc jiudged capable of the required 30 min operation subsequent to a 7.62-mm bullet impact. However, the probability or functional failure after receiving a 12.7-mm hU would be quite high unless the beating and Sear components were inordirsati'iy large. Configurption 3 is shown schematically in Fig. 416. This configuration has bee arranged to defeat 12.7-mm threats with far less weight penalty than would be incuticd by oversixing the elements of Configuration 2. The overhung mouanting of Configuration I and the straddle mounting of Configuration 2 have been couinbined in this redundant or composite system. Both pinion and gear members are supported by two conventional cylindrical roller bearings and one duplex ball bearing pair. Emergency thrust shoulders art. incorporated on the shafts adjacent to the integfal roller bearing inner raceways. Sufficient axial cekarance should be provided betweeni thec roller elements and the inner racv thrust shoulders or flanges to preclude contact under normal operation conditions (including extreme cold
I
when the light alloy housings have contracted roeitivt to the steel shafL4. H*A#ever, upon functional failure or either duplex ball bearjug. emergency axial locaition is provided by theae thrust flanges. Wse of a thiec-bearinS systcam p, .-wits total functional losit of any one bearing without seriously compromising the operating parameters of the gear meah, however, bearing alignment becomes more critical. As a result, one bearing of thc three-boaring system must be designed with greater ikiterna! -clearance than normal. The spiral bevel gear set shown in Fig. 4-16 has been enlarged slightly relativen *o the geai set shown in thea prior two configurations to derreace vulnerability to 12.7-mrn hits directly ini the gear elements. While numerous other configurations and types of bearings can be used to accomplish the same objectives, the logic used to providte inherent survivability itrnains unchanged. Simailar principitct should goivern the design of the entire dr-'ve subsystem. Their npplication, of coursc, beco~mes more involved as tae complexity of the gtearbox design increases. All shaft couplings, joints, hanger bearings or
3
ow'CONFIGURATION
D'DU:LEX BALL BEARINGS
W IT)
-12.7
*
7
-
Figure A-1,6. Tall Rotor Gear'sox
.
mrm Proof 4-21
13illow block hoirsiwigs, Warotor and intermcdiatc
4-112.1.44 FAmwgsuy Lubricadorn
lyjain gearbxec, and input/output quaill asiemblies tmust be j ined, retained, or mounted wit~h a sufificiamt numaber of redundant fasteners to preclude
1, ic preferred method of reducingl vulnerability is to assure fail-afe or emergency lmbrication in the event of total lows of thz normal luboicant sup*l This rANNhty must all'iw continued. Wae operatior for 30 min at minimum cruise power at mnission 9TSS
loss of function fiomt a single proje1cil. For well
.4prldattachment points, four fusicners ofttit suf-
`-'ic.oveve. otatins shaft joints ard ocaplings ýoften require six or morm fasteners. Frequent UKC Of -flanges, ribs and abrupt scetion changei; in castings, housings. it 1 simiiar structures providc effcti-vc 'stoppage of %.nckprapagatioia, while enhancing heat rejection to the. atmosphere. Internal ribs in Oil SUMP areas are desirable because or the ponsibility of cracking by hydraulic ram effect in the oil as well as by projectile impac. 4-1.2.11A.3 Sgif-" alig Su*4w Another design tecltirtiqw involves the use of selfsealing materiash in the gearbox oil sumlp area. The Inost efficitrnt material now available is defined as Type It ina MIL-T-5579, Thi* rubberized self-scaling compound originally was developed for fuel cells and can be fabricated to defeat either the: 7.t~2-rnm or the 12.7-mm threat. Another excellent defense material for 7.62-mmn threat is a cast urethane coating approximately 3/8 in. thick. With the latter material, the decin of the sump should be relutively silmple, as in a casting cope where the drag may 1-t witharawn without usc of breakaway core prints The cast coating contracts after pouring aad high residual comprersive stresis rcsults. This prestressed resilient coat then shrinks to close completely the hole lei's by the piercing bullet. The coating is relatively dense (2.0 lb/ft2 for the 7.62-mm throat), and also serves as an excellent beat insulator and noisc and vibration damper. Its density is such that the rurface. area to be coated should be kept to a minium to reduce the sittendant penalty on -sizing of the oil cooling system. Flat shallow oil pans and st-mps often provide the most efricieni configurations. The verious shaft seals must be designed so that a direct bit cannot cause all the oil to leak from a gearbox. This may be accomplished by using acnribbing labyrinth or slinger seals in #eries with the conventioaial contacting face- or lip-type of seat and by limiting the o-; flow rate at the inboard seal fecc to a minimum,.i Where external olcoolers and lines are used, current specifications often require the use of emergency o, :%ut-off valves to divert the oil dircctAly back to the transmission lubrication distributien system in thecevent of a cooler or line hit, thus preventing total loss of oil. One s3uch devicc is defined it. Ref. 34. 4-22
weight. Oil dams. wicks, and other itemns o e tain'n a minansum oil supply in the critical bearing areas are simple techniques to employ. Bell sind roller bearing cages may be fabricated ftom sacrificially wearing. self-lubricating composite ma~terial~s (Ref. 35) such as polyimides, Tefloni-filled Fiberglas matrices, and silver-plated, high-temnperature steel. The US Army Ballistic Research Laboratories (BRL) has clemonstrateo composit etilc gears that wear off on meshing drive gears, thus providing aform or geak. tooth lubrication (Ref. 36). Ref. 37 reports a successful applicaetion of a grease developed specilically for helicopter gear and bearing lubrication. However, the nnomal lubricants (M I .L7808 or MIL-L-23699) serve the cquc!!y important fuinctions of reducing fricticn and cooling. Ref. 38 desribes the reauiremeni that ihcrtnai cmuiiibriurn for the nrw "dtry" running condition be established to achieve 30-mmn safe operation aftet loss of the cooling oil. The equ.ilibrium can be established. only by main-K tami-ij adequate ruaning clearances and backlash in the bearings aa4d gears in the presence oC the therm. l gradients that exist in 2he "dry" condit.,on, with its alteredl frictional hWit sourcei and modified conduction, radiation, and convective hetat rejection vaths. The emergency "friction reducing" lut-ricants can be of value in austa..uing safe operation only in such a case. If a gear or bearing loses running clearance., a rapidly degenerative sequence of events results, in catastrophic failure. Loss of' operating clearance results in abnormally high he-at generation because the gear tceth and bearings operate under interferenc conditions with attendant overloads. Ti heat generation in turn produces an increase in the therinal gradients, resulting in a f'irther increase in overload and interferencc until the bearings sei;Le Or the gear teeth #, so hot the*.1 they undergo plastic failure. Specific metl..od& of preventing such occurrence arc discussed in par. 4-4.4-3. The general recomrmended design proceduic (Ref. 38? is as follows: 1. D'ouign for minimal frictional lesses commensuratc wita available manufactuiing ability. 2. Ca' culatc (frictional !osses for the "dry running" regime. An average friction coefficient 1 0,16 is suggested for the first approximation. Use ihis value with the mean values for sliding velocity, and load in the Psta meshes and bearing contact areas.
~
5
A
. ''
AL"~ 706=2O
*
3. Construct a thermal map with probable steadystate "dry running" tunparature gradients. 4. Redesign &I:gear Otbo bimfing elements to provide some clearance under the mopped gradients, Acdded clearance should be provided at high-rate frictional heating sources to accommodate transient conditions. For example, relieving clearance will not be provided by expansion of the gear cane until the increased heat generated by dry operation has beated ONe can. 5. Use materials with adequate hot hardness and frictiou properties for thermally vulnerable cornponents. 6. Provide self-lubrication cf bearings where possibin. Methods include the use of suitable cage materials or the use of appropriately located wick devimes 7. Re-caiculste bearing lives and Sear strc~aes for the increased clearance conditions occurring during operation in the normal lubrication regimc. AdJust design parameters accordingly; i.e., increase face widfts o:-pitch of glear memibei s, along with bearing capacities, as required. 44.2.143is
single main rotor configuration and are powered by either one or two 1~n..A shaft-driven, single-
In soecasesitmyb approl~riate. to employ armolt oprotect the vunml opnn.Ti design technique is the least.preferted bemause it adlds weight. increases maintenance Ptsk times, and penalizes the full-time payload. In such applications, integratl armor is prcferred over parasitic or bolt-on armor. Not only is the weight penalty slightly less with integral armor, but the pitfalls of increasing payload at the expetise of armor rrpnvai will he eliffinated. For most anplications dual-hasrdhess steel armor will be the moat efficient type to integrato because it can be rolled, welded, bolt-fastened, or integrally cast. Design of armnor installations is discussed iki detail in Chapter 14.
rotor driveshaft. Table 4-3 identifieb certain coniuaincharacteriktics for the single main rotor helicop~ters io cumr-nt Army use. iiiIt should bte viimc that accessary creasc with the sir.ý of thto helicopler. Lighti oli*4Lvation helicopters (LOH's) havo few accessory requ~reinents and 1,asibly no drive ,-dundancy. Ingr. ;nra, these helicopters may be flown safely without hydraulic booet of the flight controls, and the battery suffices for emeigency clectrical supply in the evmit of failure of th.- engine-driven gencrat.)r. Utility helicopters (UH) frequently require redundant hydraulic ptimp and electrical generator drives due to the magnitude of the rotor control loads and the increased clectrical loads attendant upon the larger amounts of inatrumeititation, electronics, and mission-essntial equil'ment. Cargo helicopters (f'H) often must havt auxiliary power unit (APU) for ground operation and checkout of electric-al and hydraulic subsystems. It is cornmon practicc to employ an indepnder~t aomssofy gearbox driven througN over-running clutches fromt both APU arnd main rotor gearbox to permit acprto rm te oe ore csr
4-1.2.2 Dii'. System Ccwsulgurstllm The specific requirementa for the drive system are dictated by the dctailasd configuration layout and vehicle requirements. AMCP 1,06-201 sets forth the evolution of an zpproved preliminaty design con. 11Suration, which will include detailed requirements for transmission subsystem power input and output drives; i.e., the speeds, powers, location, and relative orientation of these driives. Typical configuration requkirements for existing Army helico'flers are discusavd further in the paragraphs that follow. 4-.L MiS~ri Ma6l Rotor Drive tGyseaii The ma~jority of helicopiprs in current use are of the __
__
lifting-roto'r heicopter always employs antitorque reaction and thrust device to counteract main rotor driving torque and to provide yaw control for helicopter maneuverabilityr. A shaft-driven tail rotor tocatod at the aft end of the talboomn and arranged as eithtr a pusher or a tractor propeller is used most com.inonly. The tail rotor shaft is driven throu3h a 90-deg bevel gear set that ir. turn is driven from the main rotor gearbox by a long drivtsliaft or series of coninected driveshaftc. Thea power takeoff froir thc main rotor gearbox for the tail rotor is gear'ed to the main rotor drive downstream of the output sido of the freewheeling or overrunning. clutch iocated between the engine(s) and the main rotor gearbox; this arrangement permits full yaw maneuver capability during auto-.rotation or engine-out operation. Accessory drive requirements vary extensively and are dependant upon the primary vehicle mistion and helicopter size. These drives muy be miounted on the main gearbox or isolated in an accessory gearbox that driven by r shaft ftom the maiL, gaorbox. Secon-
.'qvircrrmnts
4-12.2.2 M1ihftlHU-raost Dulee Systems Multilifting-rotor helicopters have beer designeJ and test" in many configurationas - suchv,4 fere and -4-23
ý
TABLE 4-3. HIELICOPTER DRIVE SUBSYSTEMS
SINGLE MAIN ROTOR
-
TAIL ROTOR AND ACC'Y DRIVE DATA
MAIN ROTOR GEARBOX REDUCTION STAGES S SPEE11 rpm
____M.H. ENIN OTPT PU POWER., SPR hp
SPIRAL BEVEL.
T.R. SP6EO
SPEED,
PLANElARY
6,000 6,180 6,600
312 312 1,400
ACC'Y
2 1 1
NONE 1 2
AH-1 G
2,4 6.0130' 4,301
456 354 324
'R.G .6. RATIO
11
-
NONE NONE NONE
IN1 MED. G..
rm'DRIVES
SINGLE ENGINE 01+OHOH-58 UH-1I-
1
NONE NONE 1:1
0.67:14 2.35:1 2.6:1
1:1 1:1
2.59:1 2.44:1
_____
TW!N ENGINE UH-1N CH-3
6,600 18,966
CH+63
13,600
CHZ4 L
~
1,800
j2,500 j7.560
9,000j 7,S00
A NON4E NONE
1 1
2 1
324 203
2
2
185
2
2
185
4,302 %,030
j3
J
010 3,020
3
,j1.31:1 1.22:1
2.91:1 2.91:1
NOT ES: NE'~JRATOR. GE~ovs SONE ACE 550H Y PAL) LIN MAIN l..iARJibA, 4,20u rpm. FOR AND HYDRAULIC GENERATOR TACHOMETER rpm, 4,200 GEARBOX. ONE ACCESSORY PAD ON MAIN PUMP jN SERIES. 40OR 5 PADS ON MAIN a3EARBOX; 2 OR 3..,200 rpm FOR TACHOMETER GENERATOR AND 1 OR 2 HYDRAULIC PUMPS; 2.6.300 rpm OR 1 EACH 6,600 AND Z,000 rpm, DC GENERATOR, ALTERNATOR, COOLING FAN' DEPENDING ON CONF IGURATION. 2 AC GENERATOR. 8,000 ipm; 3 HYDRAULIC PUMPS; 3,700 rpm, 2 LUBE PUWPS. 2,50O AND 3.7G0 rpm;4. AND TACHOMIETER GENERATOR, 3,900 rpm. ACCESSORY G.B. POWER TAKEOFF, 6,020 rpm; SERVO Pump. 4,600 rpm. TACHUMETER GENERATOR. 4,200 rpyn. Ah2 ACCESSORY GENERATORS, 8,000rpm; 4 4YDRAULIC PUMPS, 2 4,300. 1 EACH 3,700 AND) 3,200 rpm; AUXI L IAR Y SE R 10 PULMP., 3,70U rpm. COMBINING GEARBOX APPROXIMATELY 5:1; 1 SPUR AND 2 HELICAL S'AGES 1 S'nUR AND 1 HELICAL STA.GE.
±ENGINE
aft dispowAu, laterally disposed, coaxial, and quadrilateral main rotor arrangements. All of these layouts
feature counter-rotation of even numbers of main
*
rotors to cancel the torque reactions and hence eliminate the requirement for nonlifting antitoi que device. All multirotor helicopters require rotor syr.chronization, which usually is accomplished by interconnect shafting between the individual main rotor &eaboxes, or by dual-oitput reversing reduction I;LtrinS in the cms of the coaxial confiaguration. In instance whert separate ewgines are located et each miain rotor 6tarbox, the crossshafting supplies power to each rotor for engine-out operation. In an~y instance, the intecmonnect drive is essntial to safety 4-24
of flight, requiring reliability comparable to that of the nizin rotor mast and thrust bearing. The interconnect drivz is located downstream from thc engne free-wheling clutches. The only multiliffin,' -rotor helicopter in current use by the US Army is the tandem-rotor CH-47. This helicopter fcatures twin engines of 2650 hp at 15.160 rpm. The engines are located in outboard nacelles high on either side of the aft third of the fuselage. They drive directly into 94)-deg reduction r,0_'TboXes that drive into a combining gearbox alko with 90-dg shaft angle spiral bevel geaA. T\,%oombining box is an integral part of the interconnect syr..hronizing drivc to the forward and aft rotor gearboxes.11ese
..
*,,
.
.
t
xt
.
-, ,7 -0
AMCP rotor gearboxes each feature a sih0lc spiral bevel and two planetary reduction stages with final output At 230 rpm. Thaccessori¶ arc all located at the aft main rotor gearbox and consist of oil cooler bir.wcr, two ckctrical gciucrators, and two hydraulki pumps. 4-11.23 Compomi HICellOpl Drive Systems Compound helicopters arc those that use cuxiliary propulsion devices other than the main liftng rotors in the forward flight mode. The majority of such designs have featured a single main lifting rotor, an antitorque rotor, and tither turbojet engines or shaftdriven propellers for auxiliary propulsaon. The only compound helicopter to undergo development test or Army use has been the AH-56. It was powered by a single 3450-hp engine driving directly into the main rotor gearbox. A spiral bevel gear stage, a compound planetary, and a simple planetary provide the reduction gearing for the main rotor. A spur takeoff located at the engine input to the main rotor gearbox drove a shaft running along the top of the tailboon,. This shaft drove the pusher propeller at the ead of the tailboom directly; and through a 90-deg shaft angle spiral bevel gear set also drove the antitorque rotor. Accessories were mounted at the main rotor gearbox and consise.d of two hydraulic pumps and a high-speed generator.
"4-1.3 TRANSMISSION DESIGN AND RATING
20
simply to achieve longer life of drive system corponents. Sufficient cycles will be accumulated at the 5-mirn rating during the service life of the drive subsystem to require the same bendiutp fatigue gear design, i.e., infinite life, as would be required for a continuous ratig at the same red-line limit. Although a 5-min drive system rating does not usually impair the operatianal capability of a helicopter with a typical speed-power relationship (Fig. 4-17), current Army specifications include a continuous drive sytem rating. A typical requirement would be a continuous rating of the main transmnission equal either to 120% of the power required to hover out-ofground-effect (HOGE), zero wind, at the density altitude defined by 4000-ft pressure altitude and 950F, or to 100% of the intermediate power rating of the engine(s) at sea level and 95°F. The effects of power ratings upon life, overhaul, and selection of standards are discussed in the paragraphs that follow.
I
-
L
i'
4-13.1 Power/Wfe tuteatleo, The mechanical failares of interest to the drive subsystem designer usually exhibit a definite relationship between life and power. The life-limiting failure mode of primary concern for a developed and serviceable gearbox is pitting or spa,:,iig of the gears and bearinbs (par. 4-2.1). However, the life/power relationship for this mode of failure is not reckoned with easily due to the many tacters that govern the
CHARACTERISTICS All elements, components, and subassemblies of
the transmission and drive system are subject to varying degrees of wear, abuse, fatigue, and other environmental hazards. In many instances, ,tandard components will provide acceptable performance for a given drive system design at a savings in cost, ease specially designed components. However, the designer must have a thorough understanding of the likely failure modes o; standard components (pars. 42.1 and 4-2.2) and the pertinent life-load ow lifeenviromrment reldtionships. It is customary to specify an input torque limit for
['I
1_
-
35')-
-
-
-
_00 C0- T-
I..__
. ." I5D-
, )
-
-
engine(s).
,
k
2
•5 0'•-"
a helicopter main rotor gearbox. Indicated to the pilot by a torquemctcr red-line, this limit may be lower than the sea-level-standard engie~s.._ rating of the Depending upon such factors as helicopter type and design mission, the red-line torque usually is specified as a continuous rating, or, in rare instances, a 5-min rating. A 5-min limit may be specified for •energency operation only. A time limit is imposed because sufficient cooling capacity is not available for extended operation, lubricants may be degraded, or
.41!
-
DESIGNj GROSS T
-MAX.ALT.GR
..
0o
W
.
6
90 120 AIRSPft.. ki
ISO
ISo
210
Figure 4-17. Typical Speed-Power Function 4-25
rdationship. The metal chemistries, heat treatnments, lubricants, loads, specific iliding ratios, velocities,
The Hertz-stress/life relationship varies significantly (Fig. 4-18). Each function shown results from
temperatures, geometric shape,, surface textures and roughnesses gearbox deflections, and lubricant chemistry (including water content, and other contaminants) all influence the life of the surfaces in contact, or more properly, in conjunction. It is not unusual to observe dramatic life differences between two supposedly identical gearboxes when but one of the given variables is changed by manufacturing scatt•r, operating variability, system wear, or en,ironmental factors. In a complex system of gears and mixed bearing types, it is generally acceptable to use Miner's rule of cumulative damage in a simplified form for life prediction. A representative root-mean-cube power ievl is calcu!ated from the assigned mission profiles using Eq. 4-15. The value of compressive, or Hertz, stress S,. corresponding to the RMC power or load is then calculated, and the life determined from an applicable S-N curve,
data representing a particular set of design and operating variables. The wide variance among these functions emphasizes the danger in the use of a Hertz-stress/life function without consideration of the assumptions and test conditions. Because calculated Hertz stress is an exponential function of load, little generality is lost by Rssuming exactness for the RMC life-load relationship and selecting an appropriate classic or modified stress-life function to predict the life of any particular con junction whose variablc3 are most nearly represented by the chosen function. As an example of the selection and application of an RMC power, consider the three-mission profile spectra shown in Fig. 4-19. The UH- IH and AH-IG power histograms were constructed from flight recorder data (Ref. 43). The third histogram was constructed using the fatigue spectrum supplied with a recent Army RFP for a helicopter with a mission role
350
I -AGMA 41i.02 GRADE 2 SPUR AND HEL;CAL GEARS II - GROUND AND CARBURIZED AMS 6265 SPUR GEARS (REF. 39) III - SAME AS II BUT HONED FINISH (REF. 39) IV -ASME DISCS 30% SLIP, CYM 52100, POLISHED (REF. 40) V - GROUND AND PEENED CARBURIZED AMS 6265 SPIRAL BEVEL GEARS (REF. 41) VI - BACHA ROLLERS, LINE CONTACT (REF. 42) 1 --
.__
--.
IV
III
-
V
\
I
250 21N 1V1
""16.
I.
I%
I
U.J CLi
C-2
50L
10~
103
10'
108
lip
100
LIFE, CYCLES( Figure 4-18. Gear Stres vs Life Curves 4-26
0'"1'
AMCP 706-202 ratio. Thus the stresses under the power fow the UHIH and AH-IG, respectively, will be:
30[
(S')iV
"RED LIft,
S~and
"4 .. 2
4
E]i...
. 6
A
1o2
(ScI4H.IG
I SRF
12
1
j.
SPECTRUM RED• LINE
]operating
S~~diction
X Itr and 6.5 X 10' cycles for the Eqs. 4-17 and 4-18, respectively. 11 indicates values of 1.5 X tO' and respectively. Thus, the life predicted the AH-IG (Eq. 4-18) by Curve I is 7.22 times the for Also,4-17) withisCurve I thethat life life predicted 3.0 times predicted for by the Curve UH-IHI1.(Eq.
_
.T
for the AH-IG, while from Curve 11 the ratio is only 1.67. Clearly, it is essential that a stress-life (S-N) curve used represent accurately the design and conditions if a reasonably accurate life preis to be achieved....
,The apparent large increase in life at equal values ¢omparison with cn t r th. snprat k•_rl eapr
S17 •-, SHAFT
=
proximately 2.0 stress levels of However, Curve 9.0 X 10' cycles,
REDLINE
16
161.200 psi (4-17)
(0.75)½ (200,000) - 173,200 psi (4-18) Curve I of Fig. 4-18 indicates predicted lives of ap-
14
AH-IG
4a
-
n
10
0-
(0.65)1(200.000)
-"
SHP, hp
10-
Fi;1par 4.19. S6ft Horsepower Spectra Histograms
similar to the AH-lG but powered with twin advanced technology ergines. The red-line and flight profile powers corresponding to this fatigue spectrum are taken from Fig. 4-17. The RMC powers for the three spectra are: UH-IH, 714 hp; AH-IG, 827 thp; and RFP, 1939 hp; representing 65%, 75%, and 69%, respectively, of the red-line powers for the three helicopters. However, because the sea level standard inermedip.te power ratings of the engines for the three hel.,,opters are 1400 hp, 1400 hp, and 3000 hp, respectively, the RMC powers represent 51%, 59%, and 65%, respectively, of installed engine power. On the assumption of no changes in lubrication state with advanci..g wear, the stress-life functions of Fig. 4-18 predict differences in the expected service lives of the same transmission system used in UH- I H and AHIG helicopters based upon their respective AMC powers. For purposes of comparison, assume that the red-line power corresponds to a maximum stress S, - 200,000 psi in a particular gear mesh. Because the Hertz strest in a spur gear is proportional tc the square root of the load, which in constant speed operation is proportional to the transmitted power, the stress under RMC and red-line power w' be related by the square root of the power
the straight spur gear (Curve V vs Curve lii) can be
explained best by the difference in the assumptions used in the calculation of the contact stresses. The spur gear analysis is based upon a cylindrical contact assumption wherein the ratio of peak to mean compressive stress is 4/i"or 1.27324. The sprial bevel gear analysis is based upon an ellipticu, contact assumption wherein the ratio of peak to mean cornpressive stress is 1.5. Although neither assumption is really valid, the ratio of the peak stress for the bevel gear to that for the spur is 1.178 for equal bearing or conitui arias. Ti11S izin accounts ro .. h..l... the stress difference between the two curves at a selected life of 1.3 X l(Pcycles. Additional gain can be attributed to the shot peening process that was applied to the spiral bevel sets (Curve V) but not to the otherwise comparable spur gears (Curve III). Life Rating 4-1.3.2 Tfausmluo Ovetrk The various gearboxes, driveshaft assemblies, and bearing hangers that comprise the typical drive subsystem of Army helicopters in the past may have had widely differing times between overhaul (TBO). Main rotor gearbox TBO's ranged from 500 to 1200 hr, tail rotor gearbox and bearing hanger TBO's were as high as 1600 hr. and driveshaft assembly and accessory gearbox TBO's ranged from several hundred hours to unlimited intervals based upon conditional overhaul. Specifications for next-generation US Army helicopters call for much higher (3000-4500 hr) MTBR 4-27
'
¶
without dictating TBO valuts. However, using the relationships of par. 4-1.2.1.2, application of a 2000hr TOO requires attainment of a 6002-hc MTBF to achieve the. I 00-hr MTBR (par. 4-1.2.1.2). Although this MTBF concerns only failures of sufficient importance to cause gearbox removal, it canvot be attained easily. The ultimate design goal is conditional removal without scheduled TBO levels. Achievement of this goal requires the use of reliable and thorough diagnostic techniques (par. 4-2.4.2) and failure anodes with low rates of progression so aperatioai can continue at least short-term without compromise of safety of flight. The question of a cost-eflective overhaul thne, one that balances the increased cost of overhaul due to possible extensive secondary dwtage and cornosion against the loss of residual usefuk le, is beyond the scope of this document. A coMt analysis of TOO based upon direct and indirect operating cost -A the drive subsystems of light, medium, and heavy twin-engine helicopter is reported in Ref. 44.
tion) design standards and specifications for gear tooth bending, scoring risk, case hardening practices, and gear precision clasifications amx excellent design starting points. However, experience accumulated through development and field tests will suggest further sophistications and modifications. Many useful standards and specifications are published by the Society of Automotive Engineers (SAE). The smaller size bWaring locknuts and washers are useful, but for larger bearings the SAE parts generally are too heavy. The thread specification series also is ideal for use with special beating or spline locknuti because the series includes sizes cornpEtible with standard bearing bores. The 30-deg pressure angle involute spline and scrraticn standards will suffice for most spline applications and lend themselves well to inspection wvith simple gages. In special instances, where greater precision is icquired to improve load sharing among the teeth or to im.. prove positioning or location accuracy for the mating members, a standard spline can be modified by reducing the involute profile, lead, and spacin" tolerances.
4-1.3.3 Trasssdo Stanxrds and Ratings The use of available standards in detail design is encouraged for many reasons, not the least of ,vhich is cost reduction. Available stindard3 can contribute to lower costs if it becomies umn.•',c ry to prepare special specifications; conduct qualification tests; procure special tooling; and othlZrwise compound procurement, storage, and supply activities. The standards available include military (AN, MS. NAS, AND, Federal Specifications, MIL Standards. and AMS) and commercial (AGMA, AFBMA. SAE. AISI, and ANSI). However, the limitations and ratings of standards must be thoroughly understood to prevent their misapplication, These following are some instances in which it -Is better to select a nonstandard part: i. Excess cost or nonavailability (many published military and commercial standards never have been manufactured) 2. Insufficient strength or inadequate properties (published standards for parts such as studs may not provide the required static and fatigue strength or corrosion rewistanoe) 3. Inadequate quality control for the criticality of the application (many published standards include an inspection sampling frequency that is inadequatt for critical applications), Sonic recommended uses of commercial standards are dissuaed in the paragraphs that follow, AGMA (American Gear Manufacturers Associa-
cordance with the standard AFBMA (Anti-Friction Bearing Manufacturers Association) metric envelope dimensions, using the Aircraft Bearing Engineers Committtc (ABEC) and Roller bcearing" Engineers Committee (RBEC) precision grades. De-. partures from standard envelopes may be necessary for very light series, large bore bearings; but the cornmon bore size, width, and outside diamct-r increments and tolerancs should be retained to facilitate use of standard inspection equipment by the bearing manufacturer. Cylindrical roller diameters and lengths will vary among suppliers and may not follow recommended values. However, individual rollers with one of two crown kadius or drop values are usually available from all aerospace geade suppliers. All ball bearing suppliers furnish balis In 1/32in.-diameter increments and occasional'y in I/64-in. increments. Standard grade tolerances in microinches govern sphericity; e.g.. grade 5 implies 5 pin. sphericity. Many special steels, frequently called "tool-steels", using consumable electrode vacuum remelt technology, are finding increasing use in helicopter gcarbox applications. The chemistries of these steels are identified only by AISI (American Iron and Steel Institute) specifications. It is frequently necessary to add special limits on trace elements and inclusions to thas specifications to make them comparable to some of the commonly used AMS (Aerospace Material Specifications) grades.
VV Y S.11GVGF
4-28
poeibc.e
bearings sflmowi
cIin
Qj
,
ac-
'
-.. <'
r
f
4-1.4
r
)
QUALIF'ICATION REQUIREMENTS Qualfictionequrerleut ar deflbc ~ MCP 7W6203. lPowever. there nsre a number ef qulfia tion~~~~~~~~~~~~ ~ reurmnatirins sdr~ nerlt the drive subeystemn design proems. The confldance level 1cr passing qualification tests with a mininium of redesign and retest is increased sigraificantly by rigorous attention to thes requirements during all phases of component detail design. This paragfaph treas te or:cost rquiemet. 1. Component and eaviror mental performance 3. Developmenttesting 3. Life streswntetiig tetngeude as4.e Lired su etarion testing as deaildesgn.tions tey i~et 4-1.4.1 Comaponent and Eavirorimci Many components of thc typical drive system must be qualified initially through individual twsing. Such lubrication system componei~ts as scaevnge and prs surc pumps, filters, pressure switches and transmitters, ternperaturc switches and transmitters. chip detectors, level transmitters, jcts, presiure regulators. and monitors are best defined by unecification control or source control drawings. Qualification tests are classifieL. as functional, structural, or environmental; and care mitst be exerciscd by the designer in designating applicable qualification and quality assurance requirements. Structural and endurance tests are destructive by nature and therefore arc lirnitec either to prototype qualification or to random sampling in production. Functional 'and environinental test areas may be specified as P quality assurance requirement, with the sampling ratc up to lO0%. In addition to obtaining certification of per-
be established by reducing iniet pressure to uimulate limit altitude opeation, allowing chcvks for cavitaticn as well as volumetric flow efficiency. Similarly, hermetically sealed pressure switch might be subjectcd to a wide range of vibratory frequecu.ies and amplitudes during functional or endurance testing in thc presnce of an environment involving elevated temperature and 100% relative humidity. Attention to detail of this nature can save time and duririg subsequecnt testing or field evaluation. 412D~~psu ed In the broad sense, development testing may inboth design support testing and the evaluation of' prototype hardware on bench test rigs. These funcarc an wxensieri of d-etail design whereby early confirmnation of design assumptions or errors is achieved, and noacessary modifications of the initial dtiigrs are identified. A thorough initial developmcnt test program may include: I. Static tests of' castings 2. Dceflection tests 3. Gear contact tests 4. Assembly and disasemnbly tests 5. Lube eyiAczi dcvbugjnS 6. lmnfemcntal load and efficiency tests 7. Thermal mapping tests.
appropriate to perform functional tests as a part of receiving inspection. The applicable qualification and acceptance reqL~iiemei.ts must be set forth by the designer and incorporated into the subject drawing or specification. The need for completeness and accuracy in creation of' the specification control drawing cannot be overewriptiasiz-d. If'the designer cannot identifyt or anticipate the characteristic failure mode of a component, it will be necessary to establish thorough environmental test procedures for conducting the functional or endurance qualification. Once the characteristic failure mode has been established, it is often possible to increase the effectiveness of the quality assurance t#*sting by concentrating on a particular index of perforinanc while Jirninating those test factors that reveal no useful information. For example, the critical condition for the functional test ol an oil pump may
relative locations of the maximum fiber stresses. Strain gages With suitable temperature-correectd bridges or crack wire then should be appiied at the locations of these maxima, and the casting should be loaded in increments to failure. Recorded data will demonst~ra~e compliance with stated requirements and must be correlatW. with analytical predictions so that cccurate safety-of-flight dci&.iions can be made based upon subsequent flight loan -5urvey data or when mnaterial disch'epancy iand review report action is required with respect to production hardware.
4-1.4.2.1 S-ittl Castlag Tests The designer and/or structure analyst will predict critiva' sections of tlhe castings based upon the assumed load data. A tcst fixture capable of applying and reacting these loads in a manner analogous to the intended helicopter use will be designed and employed. Stress coating and examination techniques
41A.4.2 Detflecitli Twas Deflwection tests often ar~t; used to obtain data relating to hot and cold static torque and external load deflection. These data are often required in connection with gtar tooth contacts and for verification of' spiral bevel gear development. Useful information 4-29
ao may be aquirnereaibang pla ary pear cormlitnmt. Deflection test plianot and contact rms
External fine, how, and electrical connections should be examined for proper fit and location.
K _T
data permit accurate dstrmination of planet load
sharing and analysis of cumulative tooth spacing error; adequacy of beaz.,; aleent and mounting rigidity and clamp out torque levels also may be inoutput shaft of inpat ma and be vestigated. Compatibility deflctios • equiemens wih valuted deflectiens with sal requiremetnts may be evaluated ini additiou. Multiple dial indicator gages customarily are used with incremental load opplications to obtain the re
quired deflection data. It is generally neosary to perform extensive housing modifications, usually in the form of strategically located drilled holes, to permit suitable measureicrirts. Application of red-lead paste to gear tooth members, followed by slow rotation of the drive system under the designated torquc or braking load. is used to obtain witness contact patterns. 4-1.4.2.3 Commact Tests The red-lead paste technique described *n par. 41.4.2.2 supplies information only on gear contacts. A umoie aophisucatcd coniac tmi thait al, innit* dctailed bearing study and yet does not need extensive housing rework uses copper plating and gas oxidation. For this test, contacting elements (gear pinions and inkier and outer bearing races) are flash. copper-plated (-0.0001 in. maximum thickness) and the gearbox then is assembled without oil. The assembled gearbox is mounted in a manner that simulates the helicopter installation and each shaft seal location is vented to permit air escape. The input shaft is rotated slowly to permit rolling elements to acsu.me their proonr 1vw-_tionnu and then statir tnfhume
equivalent to design rating power is applied. A reducing gas, such as HS2, is slowly bled into the gearbox, preferably through an orifice near the upper gearbox surface. All exposed copper surfaces are oxidized to a black tarnish while the Herizian contacts remain untarnished. After the gearbox is purged completely with fresh air, the unit may be disassembled for detailed examination and evaluation. All contacts should be carefully compared with debign assumptions. 4-1.4.2A Assembly and Disasembly It is essential that the designer evaluate ease of assembly and disassembly of the gearbox, suitability of standard and special tools, absence of physical interference, and opportunities for incorrect assembly. Special attention should be given to suitability of torque values specified for nuts, safetying provisions, and absence of thread galling or neizing. 4-30
4.1.4.2.5 Lubriatioa System D• r b h gp t a ih Early in the gearbox bench testing procte , attntriu tin should sh e be yc n ms. Prop-or Pe oil oiletjet distribution verisystems.
.U. I
fled. Usc of transparent (acrylic or Plexiglas) windows and covers wharever possible is helpful. Adeinternal be baffles., oil wrapers, changes incorporate should Design verified. quate oilto scavenging and intrcompartmenta venting are not uncommon.r 4-1.4.2.6 Icreumetsi Loading and Fflkldcy TeWs Immediately following lubrication system testing it is generally desirable to proceed with incremental power step testing, with disassembly and inspection taking place between each step. Intervals of 25%, 50%, 75%, 100%, and 125% of design power rating are recommended. Operation for 2 to 5 hr at each siep is desirable to achieve definitive wear-track markings at thermally stabilized conditions. Visual inspectic-i of the wear pattern: of al! gcar meehes shou!Id be mna,& after each rtep paying special note to the rate of tooth pattern fillout in order to verify use of proper initial tooth shapes. The use of black oxide on gears and bearing rings between each load step will assist in accurate visual inspection. It may be convenient to schedule efficiency measuraments simultaneoisly with load incremert testing. One rather involved but satisfactory method of accurate efficiency determination requires the external application of insulating material to the entire gearbox housing and subsccuent measurement of the oil temperature drop and flow rate across the oil cooler (Ref. 45). Power loss to the cooler PLC is
PC-FGSC,,AT
PLC
42.4
h hp
49 (4-19)
where F Gs -
oil flow rate, gpm specific weight of oil. lb/ial (a function of temperature and aeration) C, M specific heat of oil, Btu/lb-*F (a function of temperature and aeration) temperature ofT differential between oil out of transmission and oil out of cooler, deg F
With the assumption that the insulation is effectivc in preventing cooling convei n, PLC is the only power
•
(
-
loss from the transminsuion. In this cas the transalis&ionelbicienty q~, is it
1P
10,%
i
(-D
/7ý
.
where power input to transmission, hp P, Another satis'actory method for determniinng power loss is bcsed upon convection cooling and icquires the assumption that gearbox efficicney does not change with slight changes in viscosity within the range of lubricant tcmpcrature used. The exterior surface of the gearbox is gridded into approximately equal areas with centrally located temptrature sensing points. The individual areas should not exceed 5t6 in!. An oil cooler or heat exzchanger with a controllable cooling rate is employed and rair flow conditions about the gearbox are maintained as constant as possible. The test procedure requires stabilized operation at two discrette oil cooler heat extraction levels, preferably with ternperaturt. levels of the oji oui of the tramsnnission, at '.cast 50 dcg F aptart. During each of these runs the power loss to the cooler (Eq. 4-19N is measured and the temperatures of the designated' case monitoring points art recorded, along with the ambient air temperature. The increase in oil cooler heat rejection at the lower stabilized temperature coaidition is assumed equal to the decrease in convzction heat rejection from the housings into the ambient air, allowing the solution of the following simple set of equations (the primed symbols indicate cold condition): Hot:EPL - PLc + Ccs-(Y7s 2L Cod:Y
+Cc(7-
-
T,)
(4-21)
~)
(4-22)
C'L-total power loss, hp P~c. - power los&to oil cooler (Eq. 4-19), hp
S- cane convection cooling coefficient, hp/*F average of external surface temperature readings, OF TA -ambient air temperature, OF Because P;C> PcO 7"S>T7,' and XZPL and C. arc constant by definition, we havc the immediate solution: 01 (4-23) .hp/*F - PLC) (i CC Ts - r; - 5T + T)2. ' A S'S
T~)veloped
Substit~ating Cc into either of the initial hot or cold lots cquations (Eq. 4-21 o;- Eq. 4.22) will y~eld the
total power loss PI lIfthe necessary temperawrws ama measured for each tist condition. the individually calculated values of Cc may '3c averaged and a probable error computed by stan- ard statistical methods. 4-1.4.2.7 Theraid Mapping Tests Time and instrumentation capability permitting, final design modirIC46ions or the proportioning of lubrication distribution, alonj, with necessary adustzient of bearing parameter such as clearance and In-7 ternal. preload. may be accomplishe by thermal mapping. Thermocouples embedded in contact with becaring inner and outer rings and with gca blank nms or tooth fillets, for example, whoulid be used to construct a thermol mapi of the tanumimaaon. Messurement of rotating cousponrif temperatures Me quires the use oif slip rings or similar devia. Thw use of infrared photographs of opersting gearbose. also has been very effective in thermal mapping. Hot spots or excessive thermal gradients as' cause for *orrective design measures.
~
~
~
5V
~dg-
Overpower testing, sometimes referred to as weak point testing or modified stress probe. testing, is intended to yield rapid results to enable the designer to make timely charnges. Th~e purpose of this testing is to producc failures and definec failure modes and failsafe features, not to demonstrate rmliable extenided operation. However, a 100-hr failure-free overpower test at from 100 to 125% of maximum continuous power on two samples certainly would indicate that the gearbox was ready for life substantiation or qualification testing. The maximum recommended overpowei test icvei is1201-130% of normal red-line power, although in some instances 110% is used. For valid test results, the following conditions described in the paragraphs that follow should be satisfied: 1. Lubrication states should remain unchi igcd for the main power path ecu-ponents (Fig. 4-3). EHD film thickness as predicted by thc Dowson equation (Eq. 4-5) is relatively insensitive to load (125% power should reduce h values by about 4% from their 10D% power levels for an isothermal condition); however, because the temperature of the conjunction may incrrase as the 3/4 power of load, which in turn will reduce the viscosity of the typical MIL.L-7808 oil by 28%, and of the h value by 22%, a cautious evaluation is demanded. Excessive deflection must not occur. If debevel gear patterns degenerate excessively, their reduced area, cou.pled with the inci eased it%oth load, could result in doubling unit stresses ht the 4-31
-
r [
AMCP 706-202 overpower levels. The "'small-cutter" and oiher types of spiral bevel gears tend to resist pattern shift with increasing power and are good candidates for successful overpower testing. Iir well designed planetary gear reductions, it is not uncommon to find a 50% increase in uait stress for a 125% overpower test at constant spvcd. 3. The mechanical limitation of ball bearing load path constraints must not be exceeded. There should be sufficient race shoulder height and bearing mounting rigidity to retain the ball path fully at the overpower test condition. 4. Cylindrical roller bearings should have suf-ficient racetocrown) to preclude ent roller lercong crown (orra o prcreade s severe ever end lo an du eton Ssimple Hertzitan deflection. -vrpwe. The increased thermal gradients preent during Sovinrpower testing must not result in excessive bearing preloading or gear misalignment due to housing distortions. Design criteria for successful overpower testing must preclude gear toatli bending fatigue failure, case crushings, or scuffing (scoring) failure modes. Acfencatidtewear without ceromi eon of s t desin function of the gearbox for the specified test intervan is te criterion of success.
4-1.4.4 Other Life and RelabUilty Sibsbmatlatlon Testing A 200-hr qualification test is required by AMCP 706-203, and follows the tests in the preceding paragraphs. Also required are a 50-hr preflight assurance test (PFAT) and a 150-hr "must pass" qualification test in a ground test vehicle (GTV). Beyond these tests, it is frequently desirable to conduct extended bench or GTV tests to assist in the determination of initial TBO levels and to uncover failure modes not detected in previous tests. All testing in these categories is based upon spectrum loading conditions. The selected spectrum should have an RMC power level in excess of the anticipatcd flight spectrum. Because most lubrication system elements (including shaft seals) exhibit failure modes that are insensitive to power levcl, no meaningful accelerated test programs exist for the lubrication system, and its evaluation requires the accumulation of many test hours. Although the majority of lubricat;on system components will have undergone some degree of evaluation in early tests (par. 4-1.4.1), evaluation of their performance in the total system environment must await these extended time or endurance tests. 4-32
4-2 TRANSMISSIONS 4.2.1 FAILURE MODES Many competing failure modes exist simultancously in any mechanical transmission device. The modes rewognizod as dominant are often representative of the life-c' cIe phase in which the observation is made. Recognition, classification, and definition of safe operating limits are fundameptal to successful design. Failure modes may be identified as primary and secondary for ease of analysis. In one study based on component replacement at overhaul for the UH-i and CH-47 gearbox, secondary failures were shown to exceed(Ref. primary failures by at least an order of magnitude 46). Although the majority of design effort is directed toward preventing primary fai! area, the cost of drive subsystem maintenance and overhaul reflects the total of both categories. Therefore, reduction in secondary failure modes is an important objective for future design.
4-2.1.1
Prfry Faili
Modes
ae fmp odnan iderv iaiecausethosoet render a comomnnt unservicenaif because of some .self-genrated conditional occurrence other than normal wear. Cracked, broken, pitted, or spalled eluments that fail while operating at normal loads, speeds, and ,nvironmental conditions are representative of this failure category. There is a reasonable statistical level of occurrence for primary failures, perhaps on the order of 0.5%/ 1000 hr, that typifies the normal dispersion associateed with acceptable and cost-effective design practices. Failure rates in excess of this level aor considered a result of design or manufacturing deficiency. Identification and elimination of components with excessive failure rates arc the objectives of the qualification assurance testing outlined in AMCP 706-203. Properly designed and manufactured drive zystems must not exhibit catastrophic primary failure modes. It is not unreasonable to expect primary modes to be exclusively noncatastrophic. This criterion may be satisfied by inherent redundancy in load paths or load sharing, or by failure prcgression rates that arm commensurate with available built-in failure detection and die.goostic d,. vices. ,.onscientious application of classical structural analysis methods as modified by relevant test and service experience, coupled with adequate quality assurance methods, effectively will eliminate static and bending fatigue failures. However, the surface durability of loaded members such as gear teeth and
9"
,,A
706-202
roili-.4 element bearings is by no means thoroughly understood or easily preuicied. The interaction of the effocts of friction, lubrication, and wear (the modern discipline of Tribology) is the subjoct of intensive rcewm•h (Ref. 47). Drive dvsign is influenced by variables suct, as metals (hardness, microstructure, chemistry. cleanliness. residual stress), finish (routhness, lay, texture), surface treatments or coating, lubricants (base oil, viscosity, additive package), moi3turc and other contaminants, speed, slip, Hertzian stres., contact geometry, friction, and temperature. These variables, separately or in combination, may vary observed life at constant stress by a factor of 500 in conventional helicopter applications. Their combined effects also exhibit slope variations from -5 to - 12 of log.log SN curves. Because it is impossible to consider the
t
quantitative effects of all permutationa of the pertincnr paramnters deoribed in current literature, the significance of relevant test experience cannot be overemphasized. The classical stress-life equations or published S-N data must be viewed only as starting points. Table 4-4 presents useful qualitative influences of qome of the variables affecting S-N chata4,teristi.. There are many combination effects among these variables, but virtually none that result in contradiction of the indicated trends. The presence of relatively high slidc/roll ratioz and thin lubricant films is necessry for the surface pitting life to be sensitive to the additional factors shown in Tablc 4-4. Pitting or spalling generally is considered to be the result of metal fatigue due to cyclic contact stress. Under idealized conditions, the initiation of pitting occurs at a considerable distance below the
TABLE ". LIFE MODIFICATION FACTORS VARIAUt E
INCREASED LIFE
-
-_.', :
SURFACE DURABILITY
REDUCLED LIFE
OUALIFICATIONS
MLIALS
HARDNESS
RA60 -- 63 W-0
RETAINED AUSTLNITE
<10% < 5%
*
60
CAR3URIZED AMS 6260 MIQI
E2ich)
M-DUj
Ž.15%
CARBURIZCD AMS 6260
>
AISI 52100
5%
WHITE.TL lAYER
REMOVED
PRESENT
AMS 6475
CLEANLINESS
CEVM
AIR MELT
INCt UISIONS % TRACEL ELEMENTS...
RESIDUAL STRESS
COMI'RLSSI VI.
TENSILE
SURFACE TO MAXIMUIM
" ',
SHEAR DEPTH SURFAC E IINI SI TYPE
HONED. POLISHEL
GROUND
VERY IMPORTANT AT LOW VISCOSITY..
LA1. YO
3_;GrI;
TO GoLiDiNG
Vi 'i i......NI
G..
SumRACC SURFACE TREATMENT
BLACK OXIDL
BARE
LIGHT ETCH
AS MACHINED
HIGH VISCOSITY
LOW VISCOSITY
V,4 V2< 2X00 It /min
MINERAL BASE
SYNTHETIC BASE
TRLJE AT NORMAL STRLSSES
ADDITIVE IN SYNTH
ADDITIVE & MINERAL
VERY TRUE AT I OW SPEED
HIGH COEFFICIENT
LOW COEFFICIENT
PRESSURE VISCOSITYv
LOW ACIDITY
HIGHI ACIuITY
VERY IM'ORTANT THIN
LU8. FILM LUBRICANT
RAPID SURI ACE BRTEAKIN
-
COEFFICIENT, a DEGRADATION -TIME
AND
USE WATER CONILNT
LOW
HIGH
WATCH DEGRADED SYNTHI.
SPEED
HIGH
[OW
EXCEPT ROUGH SURFACES
SLIP
LOW
HIGH
POSITIVE
NEGATIVE
LOWER REL. SPEED - NEG.
FRICTION TEMPERATURE
LOW LOW
HIGH HIGH
SURFACE CONJUNCTION
GEOMETRY OF ,
HIGH
LOW
-
a-
b, a FOR ELI lPSE h AXIS 11 TO ROt V AxlS.iTO ROLL ING V
4ING
AMCP 70&,202
-
surface at the level of mayimum oithogonal shear stress, and classical theory has been dtvclopcd about these conditions. However, rtcent studizs establish that the shear stresses tend to bc located nearer to the surface. even in the presenc of very s small magniuadces of slip ( cf.48). The traction stresses imposed by sliding can raise the surfac shear stresses to within 40% of the maximum Hertzian stress (Ref. 49). These conditions lead to surface initiation of pitting or cracking that ultimately msull in the gross pitting or spalling failures ohserved in most failed mechanical comp5iients. For example, it has been established that the vast majoriqy of pitting (spalling) failures in UH-! and CH-47 helicopter gearboxes are surfaceinitiated (ReF. 46). 4-2.1.2 Secondary Failure Modes Secondary failure modes are all those modes that arc not classified as primary. By definition secondary failure modes do not contribute directly to cornFonent MTBR; however, they contribute greatly to the cost of overhaul, and in some instances they limit .......
L...-
.. rk.. .....
.:.i *; ....
o..----,
of f
primary failure. Secondary failure modes are grouped into three categories, each with a different design avoidance technique. 4-2.1.2.1 Overload Failures Components that are overloaded due to the failure of a parallel or series connected load carrying membet frequently result in secondary failure in a short time. Tandem thrust bearings or multiple planet or epicyclic gear trains are typical parallel lond-path configurations. Such components limit the progression rate of a primary failure by an automatic load reduction resulting from increased deflection or %ear material removal of !he failing primary cornponent. In such designs, the secondary load-carrying members should be analyzed under full power to insure adequate life for safe continued operation. Such analy,,es should show a minimum life of 100 hr. Series-connected secondary failures are typified by transfer of damage from one gear member to another in a train arrangement or by the upstream overload of a component duc to an bdvanced downstream failure such as a "jammed" rolling element bearing with advanced retainer or ball fractures. The static yield strength of the primary power path cornponents (gears, shafts, bearings, couplings, etc.) must be sufficient to withstand the maximum red.line power plus the incremental transient load required to fracture and break clear the rclati,.ely frangible primary failed component.
*
S~4-34
4-2.1.2.2 Debris-caused Failure Debris from a spalled tooth or bearing often enters another gcar mesh or bearing and results in suffi,:ient denting, embossing, or brinelling to initiate a secondary failure after relatively few cyclic stressings. White the da mage incurred in fears and cylindrical roller bearings is less severe than in ball bearings (due to the preferential debris entrapment of conformal contact bodies), the rate of replacement of sccondarily damaged parts at overhaul has considerable cost impact (Ref. 46). Much potential damage can be avoided by compartmentalized designs or by use of shields or baffles to protect dynamic components by deflecting and re-routing debris to catch-trap or sump areas. The objective is localization of the damage to the primary failure component. 4-2.1.2.3 Er.vironmnntally lndjccd Failures Oxidation, stress corrosion, galvanic corrosion, and aging or embrittlement fractures are examples of failures that could hecome significant to future MTBR data for helicopters with increased TBO in.tcrv,,.l
Thtsp fuihire" r,&iitt from inaderitnte atten-
tion during design or production quality control to materials, protective plating, or finish coatings. Concentric Enodes always should be used during plating of tubular shaft mcmbcr.; to secure adequate protection of internal surfaces. Special attention must be given to providing drains to eliminate trappea water at gearbox stlid, boss, and mast seal locations. Additional protective practices are recommended in par. 4-2.3.1. 4-2,2 DYNAMIC COMPONENTS The dynamic components common to all drive subsystems are: I. "oothcd poN cr transmission wheels (gears) that operate over a wide range of rolling or total velocities with moderate to relatively high sliding or slip velocities, under Hertzian stresses rarely exceeding 250.000 psi 2. Rolling element support members (bearimigs) that operate at similar total velocities, lowe; slip velocities, but generally higher Hei tzian stresses 3. Interconnecting members (shafts and couplingi) that are splined, bolted, or welded together and to gears, with external forces and moments impose:d while rotating 4. Other miscellaneous elements such as shaft seals, nuts, and locking devices. 4-2.2.1. Gears Helicopter gear design will be viewed from three aspects: limitations, analysis, and the drawing or
specirictitioii. The pirinary crairtlassis ia upon power tra.-smission g anrm ratkae than tarqu4. Sesing (kI. kind/low spood, as ipi actua=o or hoist applico'tiora) -or aommory gcarivii. The latter is diicnKWd btiefly in per. 4-5.
.f *
.
4&2.2.1.1 Gmea l~Ankeow Sucomsful helicopter Seat 4eaigns usually hovc employed counter-forma! involute spur, heiical, andi spiul bevel configurations. Somewipl~atiosnotaly te Ws~l~d WG13. haeuerofrmal circular arc hiclical tooth forms. Although somewhat superior in pcrfermanct. with respzect to skirface durability, coaformul gearv inave. inamerous configuration~al limnitat.ions, e.g., operating cantet distance is very cri~lical, and frequ;.nt!y exhibb'
precision helicepter State orauing in syntthemic tutbine lubricaiiu. The rn1~tive posittons o^ each zonal dcmaacutbn wili vary a_ a function oi the diwimers1 pitch P,4 , priusure angk 0. contact ratio, root fillet form, surfact! finish, naid tnatc~rial-procmsing cliacstoristics of each individual desi~gn. Fig. 4-20 repr seas Irtasonaibly accura': estinvmam for a standard propo'rtions, Pd - 8.5, 35 X 61 tooth seý of fula fillet foriii, ground flank, carburized AMS 6260 involute *prar gears. The variation in the three frailu'e moda rc~ationshipp when all factors ane constant except tor dra-. nsetra. pitc may tot smen in. Fig, 4-21 (Ref. 53).
r~.Auccd tooth bending fatigue strengthi. An~Aytical
101
aenot well-suited for circular arc teeth, Walous'a an
APRASE
'1
fliG
Ref. 51. ThiiR sear (frm i~i naa1tcianiriv sen&vev tatI centor distance variation unless considerail mis-
M
/S0(N
match of tooth curvature is usod betwee n pit.'un and rtear. The immediate result of this practice is a con-/
Hortzart Ftress
with an att'ndant, reduction in the thooreti"I~ surface durability of the pinion. The deflection inhecmat in the elastic reaction of the loaded tooth introducc* a ismall dcgree of slip that further reduces tie thto-Mds Fpr retical pitting endurance (par. 4.2.1.1). Conforninai I-Ars are of contiderable' ai tcrectt frota a rcserch and developmeint. vijcpoimf, but at present, dcsian mit a w'.-anin~ful d%,%ign cnscus&Aon of this configLv&-
CiNLIT II
UMFLOFTN
WPZATI(WIT P11OEILUJBRICATION)1 NOSIGCCESSFU
PT~LrVLCI
a~hcRttosp-Fim sVld:
-O
-L~
CORLTION4
tion.
I
p
SED1,
Thoe r~ataivc efliciencics of the vari'~us spur, hdli-
HIU.1O R
I02:
cal. and spiral bevel tu~oth forms wcec dincusscrl iO.)RAG par. 4-1.2. 1.1. The liig's sliding nmid ruassitant power loss and limihed load -carry iMI cap-~city of crossed cants ehminaate their usefulnesas exompa in accessory drive applications.
II
rD1GA)
eacel~ont analytical finite elanent approach to blmodicdfor three-dimensional analysis it; described in
sidcrable increase in the ina~munt
RA(G
tf
~.PITTING
LiMIT
I
4-2.2.1.2 Goar Usaly#6 'Succcssful detail involute Sear design analyxis rethree failure modes (par. 4-2.1.1) under the particL'l~r suptrating covod*tioass for each specific gear
SCRNGL-
17 10
9
appi~cation. Initial insight mnay be obtai-wd by rcviewing the regions of domni.iant distress an shown in Fig. 4-20, based also, Ref. 42. Thir, graphic rels. tionzhip rcilecto the characteristics of caw-hartlnecd
4
5
A 71 T
0IANKTR11L PI
CH Fý
3
1
n.
Figure 4-31. Graphle Relat~mloap - Fefte Mod." vs Tocda SinI -Load
4-35
:I
~
4-2.lh±.i &Oftg Waden S&reat1ial' - nThc detenuination of bendintg fatigue risk invoives th~ree stzrs: load determination, stress cvalurilion techniqut, avid definition of the properties of qitrials usedJ. The basic bendinS strmi equations, gernrally in accon.; with AGMAA practice, may bce found in Refa. 54 (spur geams), 55 (helical gcvars), and 56 (bevtl Sears), together with computer solutions written in FORTRAN symbology. Because of the increased precision w~th whith the basic Lewia equation geomeitry factouj are tretod, the use of these rcferen= is highly recommendeo. However, for a tetter underst~anding of thle relative significance of the factors of. fectirig tooth bending fatiuc, the general AGMA igidani diacussed terin by ,-. equation will be ea tern. The basic equatien for tensile stress is S1 due to tooth bending has been defined (Ref. 57) as
71
,,
111"\ /~AIK.K\ 421)ltJ "J A~
'I'
psi /portanec ~~
(4-24)
*-
(K0ý W, - WdK
where -
IP1K ki,'psi
blank type gea!,. Inasmuch as helicopter Seer have
lighte~r back-up rim and web configurationis than those used in the reference test, the dynam~ic load analyses basnd upon these methods will be coraservative. Greater precision can he, obtained by conidrg factors other than tooth d~fl&~ions by the methods. advanced in Refs. 61 and 62. Although dynamic load factors aut4 increments have been studied by numerous investigators for tlvc pzst century. the results have lCA to only slight agree:ment. Therefore, well c'orrelated test experience ie of Creat importance. The interactioai of profile modifications, tooth errors, defluction modes, spring c*TýJAants, and inertias are of such conmplcxity ta generalizecd solutions are unlike~y to be satisfactory. However, certain qualitative observntions that find geneval acceptance are:
1, The dynamic load W. increases r~entially linparl'J w/ith
%M Anmcl enai-4 (oi-irorr, n
Il.n-nf. th,. im..
of increasiog precision with speed is unden*
where face width of gtar tooch, in. F J -geoometric shape fector, dimecnsiopless misalignment factor. dimensicnlea Km, overload factor, dimensionless K0 K,-=size factor, dimenrionless A; = dyramic load factor, dimensionless 4 W diametia! pitch, in.-' W eatohlal Both tast and analysis havc confirmed that the dynmicloa *~is nor corecly xprsse asan iyamiclod vais W, mrathorrecthln exprbeing cconte s forhe, aplyng actr K tothegea I~th oad fo.treo, Eq. 4-24in shhfcor l btote. repace bytla W,. Pq.4-2 horeorc shold b relace bysolid
S,
also ii.given by Ref. 58. This evaluation is -based upon the tooth spring rate calculation taken from Ref. 59, which in t'ani presents an analysis of the extperimefltal data of ReL 6C. taken during tests of solid
(4-25)
dyamicload lb4.
Rd'. 58 shows that the dynamic load lVd eists as an incremental load duc to tooth errors and gear drii~c dynamics at operating spted, and, thereforc, is relatively unaffected by transmitted power. An ac(4piable engineering app~roach to tl~c evaltmation of
2. There is r limiting vrlue for W,, that probabl) occurs when the duration of the velocity pulse is less than one quarter the period of natural vibration of the pinion-sear sprinS-mass system. 3. The value of W,, is proportional to both spring rate and gear effective mass, Hence, it is important to consider total deflection and mass of tooth aid rim for helicopter gear designs. Fxpressions for approximate values of thc dellection rst point of n-.,sh 811AtPVL
%' *WI 3U1%JIIf
hIJU I.UL
46ASJ&rnUIL1LPhI UG
61U13IVE
by Refs. 61 and 62. For tHin-rim helicopter gears (back-up rim thickness of approximately one tooth whole depth) with approximatcly 40 or more tneth, the defection can be as much as twice the value for a rim. The limiting deflection of a spur tooth at the momen't of mesh impact is approximately 10% that of a helical toothi, reflecting a spring rate an order of magnitude higher. As a result, the limiting
dynamnic load on the spuir gear will be about three
timcs grtater than that on the hclical gear. This alaiagy also holds for comparing w.taight and spiral bce~el gears. In mist cases the torque capacity of a hardened precision helicopter Sear set will be limited by surfNoe duraaility rathar than by tooth breakagc resuiting from dynamic loads. Many failures formerly attributed to dynamic loads have been recognized recently &&resulting from rcsonant or vibratory conditions that may occur at h;gh speeds (par. 4-2.4.1).
¶
7)AMP
7, 4
Determination of the goam tooth load W, is relatively strvg~htforward. For a first approximation itius NWt
,b
IC
3. Elatic deflecaioas of sAMUa, gear wabs and rims, and support bearings. The generally accepted rel-ationa~shi for the misaligned facto&K. for spur and hdlita gears ~rc
(4-26) DPN
K.
we:torque, ltu-in. N W M number of driven gecars in mesh with the driver, dimensionless For precibe calculations the pitch radius P,,/2 should be replaced by the radius to the highest point of single tooth loading (HFSTL) an determined by the profile contact rotio' for spur gears or by other con*Aiderations for helical and bevel gcars. These considerations are discussed further in a subsequent paragraph in conjunction with discussion of the geometric shape factoi J. The overload factor K. is included in Eqs. 4-24 and 4-25 to account for the torque pulsationi waveform or
K. US~ F' di-esa (/)'
K
~
VMSW-=tc; toD asac..
I&HAis;OuK53535@
horizontally opposed aircraft en-
sine K,- 1.25 - gear set adjacent to high angle Hooke's joint installation K. 1.15S - gear set adjacent to typical tailin rotor drive-shaft K,, - 1.2 - third-stage gearing ixi six-cyli 'nder reciprocating engine applicat.-On K,, - 1.0 - turbine engine speed teduction Sear drive. The mibalignmcnt factor K., in the tooth bending stress equations take: into account the lengthwise or axial load distribution on the face of the loadod g", mesh. Three primmy~ sources, which generally are aaditivc, contribume to misalignment: 1. Initial misalignment due to manufacturing inaccuracy or dcflectdc axes of rotation due to gearset load, external load, or thermal gradient 2. Toeth lead slope deviations due to inaccuracy in gear manufacture
, for F
.,S >F (4-27) '
'
W,' -K,,W, +W, lb
(4-.29)
Empirical expressions for Fm are
r
"a
result from the power source, the driven mecmber, or the response of the elastic drive subsystemn itself. Because the bending stress S, must be wvithin the fatigue endurance limit of the material, K. should not be used to account for occasional overloads of low cumulative cyclic duration unless the design life itself is relativey low; i.e., less than 10' cycles. SomeJ examples of measured K. values are K,, - 2.0 - first gear drive from gix-cylinder, rociprocaiti a&,four-stroke cycle,
(4-27)
whore F -~ face width of gear tooth. in. F., - average value of effective face width F., for given loading condition, in. The ave.rage face width F. in those equations is that width which con be considered 1.o remain in contact under &nef fezinive tooth load W, where, from Eq. 4-25
for roughness of the transmitted power, and hence is Q .S.Im..u
frFor9 d'less,
- 2F -
[-
-1w~ 1/2 ' J
(4-30)
in
-eL'J
for spur gears, and for helical gears F. = 2
Fe K4
1' /2'~ -
(4-31)
where e - pitch plane misalignment (net), in./in. G -lengthwise tooth stiffness constant, p: k
= enntar-t line- incdi-istinn fskrtnr (Re~f
P6
dimensionless = bas pitch, in. - total transverse length of lint of action.
Z
All~
The value of the tooth stiffness constant G in thewe equations usually is 106:! G :5 2.5 X !10'. The correct value for Km for "poikit-contacts" such as those in spiral bevel gears may be considerably less than the values for spur or helical gears due to the conformal axial curvature and curvature mismatch used to localize the contact pattern within the tooth boundaries. It is common practice to use. a value of K., - 1.1 for aircraft spiral bevel gears. This low value is particularly jusuticid for straddle mountings used in conjunction with so-called -small-cutter tooth developments". Use of cutter (and grind-wheel) diameters equal to the mean coite distance at low helix angles and equal to less than L% ice the cone distance at high helix aigles is an example that would 4-37
*
meet this criterion. In Seneral, the contact pattare
shifts toward the heel of the tooth as the lead increases when too large a cutter is used, while a shift toward the toe results from use of an excessively small cutter diamvter. Thc correct diameter will result in approximately equal pattein spreading toward bothtoe nd hel.The bthe stoeres hevel.ainprino 42 of the expression (Pd/F) (AJJ). ptchdiide byfac with P8 F) Thediaetrl defines the physical size and hence the basic strength of the gear tooth. K,is a size factor to account for the phenomenon that larger components may not exhibit fatigue endurance stress levels equal to those for smaller cornpjonents. It is grouped with other term.,; in the equation foi gear tooth stress S, so that resdy compAri.aon may be made with the basic meatenial rllowable stress. For spiral bevcl gear applications Ref. 56 recoinmends for Pd !•16 the use of K, - 2Pj"' 5 ,dirnanisionli and or 16whee foor Pdor I>
an
K. 1.0, dimensionren*
(4-32)
A;~claot viitrndcd AMS 6473 it is appropriate to usc K, - I K, -0.5 + 4.S/D, K. 2
-irosols
dinniols V Y A>
o
(4Yi (434
or aco, ieninls iesols atr -fr = stress concentration factor (Ref. 01). dimens~onleas Pý= contact ratio factor, dimensionless
for helical gear; d'css
(4-37)
16ones (4-33)
form infiatnfator, dimensionless
C, -
for D Ž_12 for I > > D 3 Or DPS 3P
(4-35)
geometric shape factor J is uscW to atxount for thet,hape, of the cantileverv-d gczr-tooth beam; and it iflCa'.-si.A e iifluienucc of atvss concentration, load sharing, and the modified Lewis fowon factor. The reom dd uainfr.1r: for spur gear,
Kh~
However, for spur and helical gcars for which 6 :5IPd: 11t, Ref. 54 suggests K, - I because wc'i data have produced only a slight strength difference attributable to size. It should be emphasizcd thut the values for K_ given by Eqs. 4-32 and 4-33 are based upon the assumptions that the ratio o1 case depth to tooth thickness remains essentially consiarn, for casehardened gear teet~h, and that the characteristics of casr. and core material and residua: stress fields arc unaffected by size. Although the letter condition can be achieved in the practical sense ri-ver a fairly wide range fc~r case-carburized materials, it cannot be satisficd for nitrided materials. For materials such as AISI 4340, AMS 6470, and AMS 6475 *latare to be case-hardened by nitridirig, a correction liecomes necessary as 'the diametral pitch increases. Whi~c there are few published data in this area, reascnable corectonsforniude4AIS 430 o AliS 470for various values of pitch diameter D,, are K, - I K,- 12/I),, K, - 4 f 4-38-
for D ýt 9 for 9 > D > 3 for D,:5
(4-34)
WeIX gngle.dcg
and for bevel gear, j
Y~ip
where Yt
R,
les 'ls
(-8 (-8
=foini facicr, dinicasionlecss
-di--tuance fron; pitch circle to point of locad applic.ation, ýn. F, - efrouivt face widtn, in. inertia factor, dimensionless K, R -mean transverse pitch radius, in. Pd transvermu diametral pitch (measured at large end of bevel gear), in. -' P, - meun transverse diamietral pitch, in.-' Eq. 4-39 is taken from a Gleason standard (Ref. 64) and is pres& ted primarily for discussion of the pertinent parameters. As mentioned previously, a more thorough and accurate evaluation can be obtained with the computer program oif Ref. 56. The modified Lewis form factor appears as Y, Y,, and Y5 in Eqs. 4-36, 4-37, and 4-38, respctiveiy. This factor is based upon bending stress calculation for a parabolic cantilever beam inscribed within the involute tooth form with the point or tangency between
AMCP 70r2021 the parabola .. 'i the involute. and the: point of load from cantilever plate bendiag theory as presented in application being the significant factors governing Ref. 65. the strese. This form factor redue the extreme fiber tensile stress with a compression component of the CA d'leus(4-41) tooth normal load. The load application point for the 1 spur Sewr factor Y always should be taken at the calculated HPSTL and At the tip of the involute gear profile forth helical Seer factor Y,. Th~i where determination of the load point for the bevel Star fac v -hlcltohla ln nlnto nl br ~Crtan ~ ase Uumpi~n Upn fl~flfl~Tan-' sino~tanV,), deg the load contact pattern geomfti of the bevewl tooth.-nomlpesrage hic or)d# -Thegrcat difference between the assumptions of Ref normal totpires lrangle (ei'ga) deg. 64 and the updatod version of Ref. 56 accounts for a -g ot prlage ~ The inertia factor Ki in Eq. 4-38 accounts for a considerable change in the calculated stress. Use of reduced contact ratio. For in0 > 2.0. Ki 1.0 and for the modified Lewis form factore in conjunction with a stress concentration factor K1 has proven to be as m~, < 2.0, K, -s2.0/mn.. Eqs. 4-36 through 4-38 are of assistance in accurate as any method known for involute Sear the stress at the location assumed to be the evaluating pn~iwrc the range of within calculation tooth stress weak point of Star teeth. Test results often indicatt. angles 14.5 deg < O< 25 dt~g. Howevcr, a significant however. that failures originate not in the fillet near degrcc of inaccracy may occL~r cutside this range as the involute flank, but rather deeper toward the root *,ell as for internal Sear tooth forms and for or higher on the tooth flank. maximum fillet re4ius configurations. In the former case, the crack propagation is freThe influence of the stress configuration due to the quently downward througi' the rim rather than in an relati'.c fillet radius and load point location is ac. .%.a-'insul .. 2; .1 tau ý J~ ~ 4'..i c.. s n t~ . of the ceptable failure mode because a large section photostress work of Doln and Broghamner. It should gear rather than a single tooth tends to break off. be notedA that an C c.,tv'c Kf is inctuded in the vabies for A inE~. ~ Diakage high on the flank may result from the use The effective ifid apportionment due to loadOt" 1i ot nargdbs 'hls-hn slarn~ t~~shiag mon eet isacconte fo in optiniun blcAd between the fillet and the flank. A lo t'-kuently arises when the back-up rim is thc conza,:i ratio :7acti.i.. c,, F'o ýgar Sears fro wh~ich theproilec~itac raio~. ao,~ I~c..iisc ll too N J,~~' i,, case the rim bending stress (due to about the rim neutral axis) cai i be niv v-nunt A To&.' tooth mn 5 2.0 For HKETL. the upo~i !vised 7i~r calculations cantilever tooth [*ending -,a~ assumed ~ta double of poipit ~iir.ct h t P wh-i •5 3.0. stress. Decauts rim curvature and web resistance toot asui~c~. coti~cor d~miatio ofti~ enter ilflu ti'~s nnalysis, it is incorrect to considcr only geosnelric ~ ~1~ 'rh~lnast~~ai~~rm ~ ~ ~ ~ ~~th n~o.fco ihicknes&. However, for spur and iow-eiciixorfaceeonact ~io ~ a.~.~ accuntd fo by angle helical. gcs'rs of about 40 teeth, a rim thickness &ii~i4 equal te the tooth~ depth is generally accepted as aide49) quate. Slighilj lesser v.-uos may be usrd for gears ~ *9 ~ us~ ý4,ý ''mnainlos 2(-3 with fe~verwe~thile greater values may he needed for gc.5r% wim~ mr4or than 40 teeth. The existence of Whbere high thb us k-wls m~ helical or bevel gears will cornp- oit!iýru~tiu pitch (helivalgc.ar). in, For ~ plicate tht: avo~.dyris. Considerable web and rim rcin. for~imýnt zi. n=&itary to dcvelop full tooth strength clb.the modified cvuasav r.tio in., Frspi&l which is a rmo-:ei anmsqiiare (ruzs) mitmmation of the po~intol . ihr~'t-mpnt prabel v..tivv profik. i&d fax.contact ratio, h; used. When beas th rcdniaaye r wo-dimen2.ý,m - 1.but for mi.> 2.0 sional, th.iv' att not suitable for definition of maxima for h tr-ei ial stress field. Computerized methods of fir.ite element analysis are expected eventually to af, d'lmq (4..40) .. rn~ ~ f-3. d accurate solutions for these conditions - 4Y' MI + 2 %rn' .4 Additie'iul prevition of high-speed gearing 4nI~lyThe CA, factor used fis Eq. 4-37 liccounLw for the ses m~ay t--tbtaincd by inclusion of the hoop stress Sh in V~ic geali rim due to the centrifugal acceleration. A inclination of the load contac line and is 4.r.~ed -a
.
A
,
.
.
p
-
-
conservative value fow this steady stress (at a con-
limit data and an applicable value for u, may be oh-
stant speed) may be taken as
tained from an R. R. Moore rotating beam specimen test. Extreme care must be taken to dnplicate every metallurgical and manufacturing characteristic of a
-. 910-p(nD)
'.
"7
'
psi
(4-42)
where 7.095 X 10-6 - dimensional constant p = material density, lb/in.3 "n Sgear speed, rpm D3, - gear root diameter, in. The oscillatory stress due to bending S, may be combined with the hoop stress S, by use of a modified Goodman diagram constructed on the basis of the material properties of the specific gear. The modified Goodman diagram can be used to account for the reduced allowable bending limit for an idler gear application in which the calculated stress is fully reversing. Use of the diagram is discussed in detail in Reft 54. A maximum safe working value of tooth stress S,.. due to bending can be determined as , ~,+K, (4-43) S allowable endurance limit stress, psi K, lf acodmesols K, - temperature factor, dimensionless Kr - reliability factor, dimcnsionless The life factor K, is assumed as unity for all Sapplications designed for infinite life, i.c., greater than 101 cycles. All Army helicopter power gearing designs must meet this criterion. The temperature factor K, is taken as unity provided the gear blank operating temperature is blow the hardress draw caett - . . ....' ...-. , hi..s 6-r;.,io.. must be satisfied by all Army helicopter power gearing designs. The ieliability factor K, effectively is a factor of safety that is used when the statistical confidence and reliability (test data scatter) are unknown for a given mean value of the eriduranco limit stress, In such casts, a value of 3.0 is recommended for K, When the allowabl endurance limit S,, is known for the specified reliability level, K, - 1.0 S., always should be chosen to rmect the desired reliability for the design application. A genS'rlly recognized safe design practice for helicopter geaing is to select S. as the value 3 standard daviatior4 (3o) below the mean endurance limit demonstrated by teat. The value of the standard deviation a, as well as the mean endurance limit, varies greatly with material, heat treatment practices,! manufacturing variability, and the quality control level exercised in final component acceptance inspection and nondestructive test and evaluation methods. Endurance -
*
2 ,
"4-40
gear itself or the data are useless. However, the value of e from these tests invariably will be smaller than that for the more complex production gear. Refs. 66, 67, and 68 present the results of indepenjdent test programs conducted to determine an accurate mean endurance limit for carburized AMS 6265 gears. The resultant S., values vary from 160,000 to 210,000 psi. Relative strength data for many materials and process as treated in single tooth (pulsor) machiqes over a 15-yr period may be found in Ref. 69. The variability of a and S,, with material and process is shown in Fig. 4-22, extracted from Ref. 41. None of the test gears were shot peened since this proess could ;..ve masked the inherent differences in the materials and processes. The endurance limit is, of course, merely one of several factors that must be assessed by the designer in making his gear material selection. Aside from the obvious criteria of fainiliarity, confidence level in hmanulacturing and process coatrol, cost, temperature environment, and size, the crack propagation characteristics must be satisfactory. Although the safe design stress for one candidate material may be far higher than that of another, it also must be assumed that ballistic damage or secondary damage will occur in any critical gear mesh of an Army hellcopter. Such damage could result in an impact-type overload that could cause a through.-hardened, hightensile-strength material to shatter and fail instan_!y The degree of ductility provided by a core structure of lower hardness (commonly core hardness 20 points Rockwell C lowe- than the case) is often sufficient to provide a safe failure mode with a relatively low rate of crack propagation. When this technique cannot be employed, the use of lower hardness gradients together with geometric crack stoppers may suffice. Ref. 70 is a useful primer on the fracture mechanics for the gear designer. The use of properly controlled shot peening in the gear fillkt area often can reduce scatter (smaller a) and in some instances can increase the design allowable S., by 13 to 25% in carburized AMS 6265 gears. Practices vary widely with regard to peening of gear tooth faces. Some specifications require masking the faces during peening, others require removal of the effects of peening by flank grinding or honing. and still others allow peening of the running surfaces. Peening requirements are tailored to the gear application, but certain generalities may be stated: I. Peen in accordance with MIL-S-13165.
w
.
-
r
C - CARBURIZED N - NITRIDED H - THROUGH HARDENED G - GROUND FILLETS P- PROTUBERANCE HOBBED R - HIGH RETAINED AUSTENITE Z - ZIRCONIUM GRIT BLAST
0J MEAN ENDURANCE LIMIT C:' -3,y ENDURANCE LIMIT (SAFE DESIGN LEVEL) 2O O , -0 •
=-U--
-• .-.
_
.ct
-)
Q
C-,^ C3-
•
'..
LU
F
__I
C) -J
MATERIAL TYPE AND PROCESS Figure 4-22. Single Tooth Pulsor (car Fatigue Test Results
2. Use cast iron, -tci, or cut steel wire shot of diameters no larger than one-half the smallest fillet -radius. 3. Use several hundred percent coverage, • 'r'a,•, ,a 4. Dc.. [iii ..ri.if. i.,. *L i"..-r Lt,'• b',-L|IauaWs .i im iw v e Wll
',
Ao,-
positioned to simulate the exact surface location to be peened. 5. Never exceed 0.016A intensities, 6. Do not peen very hard (> R, 64) or brittle surfacs such as nitrided AMS 6470, AMS 6475, or Ssimilar materials. However, cleaning and abrasive grit or glass shot at intensities up to 0.01IN is permissible. 4-2.1.2.2 Scoriag Failure When two gear teeth slide togethci under load, there is considerable heat generation in the loca:ized conjunction even in the preusnce of a lubricant- When the rate of heat buildup exceeds the rate of heat transfer away from the conjunction, the resultant temperature rise acts to reduce the lubricant viscosity, thus reducing the thickness of the separating oil film. In the full EHD lubrication zone, the temperature rise to the friction; however, as th: film
\_*4s
rve -orl
_44
thins and conditions change te the traniitional or boundary hubrication regime, the friction generally increascs. Conscqbeitly, the unstable condition may be created that eventually will result in harsh metal*. *•_-oW ii. * ',t.% ilvu
m,
o4*
a, s,.,rLr ... en-re - ...-*,. •.
,
demnit y suf--....
ficient to "melt" and smear a twin lyer of the 2ear tooth surface. This smearing condition is referred to as scoring or scuffing. Although the physics of the phenomenon rcmains the source of much debate and intenive rmscarch, ceitain ficwors are Lnderstood sufficiently well to permit engineering design that will minimize scoring risk for a given set of operating oonditions. There are five fundamenta& approaches the designer must consider: 1. Selection of the proper diametral pitch, taking carc to balance bending strength against scoring (Fig. 4-21). Higher speeds call for the use of finer pitches. It always should be possible to select a pitch range in which putting enduranc is the life-limiting failur. mode. 2. Use of a contact ratio sufficient to insure load sharing by at least two pairs of muting te-th in the higher sliding velocity ranges of the tooth contact. The profile contact ratio for straight spur gears
operating in the velocity ranges of scoring sensitivity never should be less than 1.65. a value that pcrmits two-tooth load sharing in the first and last thirds Of meshing contact. The profile contact ratio may be reduced for hliefcal a-id spiral bevel getass i,,ea thrc d face contact ratio is suffcient to assure a total developed contact ratio of 2.0 or greater. Reliable achievement of these contact ratios requires accurecy of tooth spacing, profile slope, and lead slope. 3. Seleiction of tooth numbers to insure hunting tooth action. In general, this requires that the number of teetl. in any two mrshing gears be relatively prime; i.e., that theme be no comimon factors. An indication of the significanrcý LA this requirement isgiven by Fig. 4-23, which illustrate* the difference in scoring load limit between Star-synchronized and separate motor driven 4.0 in. diameter test disc operating in MIL-L780 oil (Ref. 71). 4. Modification of the involute profile to compensate for deflection of the teeth under load-deflected load so that the loading ors the entering and leaving tooth pair contacts varies susoothly rafthr than in a * step-functon. This not only minimizes the transinitted I;;-- carried att the elidin velocity cU
The most satisfactory technique cur rently available for calculating gear scoring risk Is based upon the critical temperature hypothesis (Ref. 72). This hypethesia suggests that for every oil-meta combination, there exists a cr~tical constant conjunction temperature at or above which surface scoring occurs. Application of the co~icepi requires determiratior, of appropriate values of the critical temperature T, and the temperature of the conjunction T7+ A T with 7, the initial temperature of the oil-mesh interlace as it enters mesh and A T being the temperature rise during the meshing cyckc. The critical temperature hypothesis implies that the limiting or failu~re load Wf is related to certain design variables. The specific rzlationship has be=s reduced to standard geur terestinology and published by AGMA (Ref. 73). Aver4. values for thermal conductivity, specific heat, dentity, and an assumed constant coefficient of friction of f - 0.06 have been incorpoxatuid into the empirical questions for the temperature rise A T and the scoring geometry factor 2, ,\
ij
but reduces the dynamic ioad increment as wel!.
5. Provision of adequate, uniformly distributed oil flow to the entire tooth face width. The dual function of cooling and lubrication Ii best served by use of bo'.h in-mesh and out-of-mesh oil jets. When both jets cannot be employed (either to minimize windage losses or because of marginal pumnp supply), it usually is best to retain the in-mesh jet for high-speed gearing and the out-of-mesh jet for low-speed geapryns.
(4-44) 5OSJ\) /J(-4
where W'-ttlcfcieprtohtal F, - a effective gaer tooth, i oad lb S - effectiveface width,pin. S - rinis surfac frispin - scrn speed,- rpm r dmniols -soiggoer atr iesols
(4-45)
ANN 626eDISKS 14O d,a,0.0175
~~hINLE
1
~ j i;
T OILTCWP-Iq I
V411'
CI---. UIISVIHqOMIZiO 90~ SYIICHRONIZEU O
where
--
N
P1
N
--
r T
Z7
- number of teeth in pinion -
number of teeth in gear diameitral pitch, in. radiu~s of curvature of gear tooth, in. radius of curvature of pinion tooth, in. transverse operating pressure angle,
deg
I using Eqs. 4-44 and 4-45, T, is assumed as equal to V,-V'. 11 mm the oil inlet temperature. However, considerable error can be introduced through the actual values off Figur 4-23. Scuffivig Load is Sliditig Velodly and T, because the friction coefficient is often lower Sym~eakd ad UsyncroasedDimthan 0.06 and the pinion-gear surface temperatures 51200300C
J=
L-4-42
Usnhole
ie
x
1)P70-0 often are higher than those at th. oil inlet. However, because these two errors are opposite ip offcct, sufficient cancellation occurs to rendcic the equatio~is acoeptable for estimation purposes. Ref. 73 defines the critical temperature T, with respect to scoring risk, rating T, - 5W)OF as a high risk. 300*F as a medium risk, and lesser temperatures as low risks. Theiefore, the value of the conjunction temperature T, + AT under design load conditions should be less than the, value of T, assc vaed with an acceptable level of risk, There is no accepted or inherently accurate method for calculating T's, although measurements of typical 0 helicopter pinions have shown values lO0F greater than oil inlet temperatures. Although it is well known (Refs. 74 and 75) that above a certain critical speed the scoring load of a given gear set will increase, the AGMA equation does not reflect this consideration since no speed term other than sliding velocity was used in the deve~opment of Eqs. 4-44 and 4-45. An improved calculation ma~hod uses speeddependent friction coefficients (Ref. 76) combined with the effects of tooth load sharing. The method for dlga, ::..,.
e
uic
foa,,...,sou
ZjC
/
~
5w~~'
kq"-')
high risk, but 7, = 400*F as a medium risk thus would be a suitable classificakion for carburized AMS 6265 gears operating in lAIL-L-7808 or MIL-L-23699 lubricantls. The preceding analyses do not adequately accourst for certain factors that are. known to influence the calculated temperature rise AT awtd the truc 7:. for synthetic lubricants. Among thesc factors are: 1. The diffcreaccs in friction and wedr additive ef-
-~~~-
i
*i
p.,*~
r
Feflfit
ficiency between MIL-L.-7808 and MII..-L-23699 oiia
...
a, les
6. Calculate a value of W,' adjusted to account for load sharing at multiple tooth contact points, including effects of profile modifications. 7. Calculate AT and conjunction temperature T, + AT at each point, Proper load sharing distribution in the two-tooth contact zone must be provided for by involute profile modifications. There are many techniquts in use for calculation of these modifications, most of them based upon the practices recommended by Ref. 77. Even though the calculated values are slightly low for thin-rim helicopter gearing, the tooth defiections obtained by the methods of Ref. 78 should be used for profile modification technique. The first point of contact (pinion dedcrndumn with gear tip) is the most critical with respect to the overload effects of tooth \spacinS errors, and produces the higher absolute
.'the
*\
V\.0/
a* value*of TI *
L'M
1. Subdivide the active tooth profile into at least 20 equally spaced points, 2. Calculater andr at each point. 3. Calculate 4at cthpoint, 4. Calculate f at each point by method shown in Ref. 76 or using suitablt; empirical data. 5. Replace 2., in Eq. 4-44 with Z,-, a modified sv'ring geometry factor.
-'
values of the friction coefficient (Fig~s. 4-6 and 4-7) Therefore, to achievc the best proffic modification for scoring risk roduction, the prccedinig caltulations should be slightly biased to increase the Miaterial removal at first ploint of contact while decreasing the removal a,. the last point of contact. A 20-30% bias shift is generally satisfactory. If practic~able reduclions in scorinig risk are to be obtained throught involute modification, profile slope tolerances must be held bttween :k 0.0001 and 1 0.0002 in. for the modified zones and tooth-to-tooth spacing accuracies of 0.0002-0.0003 in. must bu achieved. Adequacy of the ctlculated design values must be confirm-A dtiritig initial gearbox bench testing. Prop~r profilk, modifications for helicopter applicotions must reveal full visuni profile contact throughout the rfaige 50-75% of the red-line power; if less than full contact is achicved, the rtsultant loss of contact ratio at nor-mal cruise power may cause excewskvely rough and noisy cptration with an attendant reduction in p~itting life When the level of sophistication desccibed is used in the calculation. of T).+ AT, together with precision in manufa~cture. the risk evaluations of T. 0
texture 3. The influence o! EHI) behavior as a function os' te erteanvloiy When tests are conducted under closely controlled conditions wherein the friction coefficient f, the initial temperature T,, and the EHD parameters are known with accuracy, it has been reported that the assumption that T,. is constant actually is invalid. Ref. 20 shows a semi-log correlation between T,. and a dimensionless El-D parameter tp which depends upon the initial viscosity p, the sliding and total velocities V, and V7, respectively; pitch radius R; and compressive (hertz) stress at failure Sc,. In this correlation the value of T, for well heat treated, low retained austenite, case carburized AMS 6265 operating in MIL-L-7808 drops from &.bout 600'F to about 430*F when the value of t increases by a factor of 101 (from i - 10- to - &~.Values of T, are approximately 100 deg F less for lower quality (with
0!
-
~
high retained ausmteit) mas carburizd A MS 6265 op~uieing in the saicie lubricant (Rqf. 71). Tbo reuctioi n.V T, with anl kincrea In ifis due to One 9omplex interaction of V, and Vy, For conlstanlt V7 T, faho 6harply and then levels off as V, is incr=4.d while T, increases eapnential ihV wimm V, is held constant. Because the ratio /VI1 . is constant for a given gear design, these opposite ef fecta tend to canWeleah other over common ranges of gSar operating speeds and loads, producing a relatively constant value of T,. When actual friction data for t given lubricantmetal combination are not availsble, the trends shown in Fig. 4-24 (Ref. 71) ane helpful in design mvKiew and evaluation, From a practical viewpoint, when overpower tests * show that scoring risk is marginal, the problem may be eliminated by such relatively minor remedial actionsacB 1. Improving the run-in cycle by i±. glonger runs ;A~t increased load and reduced speed to rewict tht operating surface finishes 2. Reducing the manufactured surface roughness through better grinding piracice or the use of Scar tooth honing where possible 3, Reducing the value of T, through increase lubrkicat flow or cooler lubricant s;Apply. When such measures prove inadequate, the lubricant and the metallurgical microstructure should be evaluated. If neither can be improved, it may be * possible to improve the involute modification or profile and the tooth epacing error, -
4-2.2.1.2.3 Pftrimg Fallure There are several pitting failure modes that in their advancad states produce the same and result: extensivc spalling and tooth fracture. Only three of these modes are relevantto the type of gearin used In modern helicopter drive subsystems. They nay be classified as case failure, classic or pitch-line fatigue, and wear-initiated failure. 4-2.2.1.2-3.1 Case Failre This mode results simply from inadequatic depth of case to support the opeating load. It way be avoiodd by adjustment of either unit load or oas depth to obtain a ratio of subsurface shea stress to shAm yield strength in excess of a particular critical value. Ref. 79 recommends that a value of 0.55 for this ratio not be exceeded; however, for high-quality helicopter ý taring transient operationc at values between 0.55 and 1.0 should not result in failure. Extended operation above the critical ratio will cause subsurface cracks to occur near or in the case-core transition area as a result of the repetitive subsurface shear. qtre-citins. Them~ suhxurfxne cracks scion spread to the tooth profile surface and generally result in numerous brittle longitudinal fracture in the general area of the single-tooth contact zone. Total mutilation of the toth profile then results from only a few additional cycles of load application. The variation of subsurface shear stress with depth may be calculated in 4astrakightforward manner. The magnitude of thie subsurface shear for a given depth is a function of S, and the Hcrtzian contact band semiwidth b.The calculation should be made for the lowest point of single tooth co'ntact (LPSTC) on the
_____________
_;.in
AM4020DU 1considerably 0U'-IK)L( OIL QAMIL- S¶RAOIGTMKL
L
-V' - V, $250iH
a -'
~
-~manner:
sur.Eace
then..
-la-,
.nmember
6.
weaker) becauscr this produces the
0
maximum~value of S,. S, may be calculated in accordance with the methods shown in par. 42.2.1.2.3.2. The effective toota load W,' and the radii of curvature shouald be adjusted for the LPSTC. The ~Hertzian semiwidth b is relatod to S, in the following
-
-(2.30
-b
j
-~~ ____
VI -V2.f Itabl
Figre 4-24. Scufling Loand vs lubricant -plotted Ussmycbtosihed Discs '4-44
invlu.,a
x 10-1) (7t L)S, . in.- (4-47)
wh!5re r~ and r are as defined pre'iowjly, the radius of curvature of the'pinion and tar- tooth, respective4-5 nlxt should bc use to cak &late values of shear stress S ait 12 depthis. Thate alues then may be along with the allowable st.resz as shown in Fig. 4-25. The allowable values show n are 55% of the
(
-J
II
-
i
TABLE 4.5. SHEAR STRESS VS DEPTH DPTn.SHEAR STFESS -
shecar yield strength, as a function of hardness at the given depth. An approximate relationship betweon to shear yield stnis is shown in Fig. 4-26. allowable values of shear stress near the sur,ace
DEPTH. ohardness
C, x b F
VALUE
C, xThe
OF C1
........
_..
0.05 0.10 0.25 0.33
1
are omitted because of the large residual comprossive stress field normally in existence therem. iB.ause this residual field will reduce the effects of tl'e imposed subsurface shear stresse, this region is not critical to the analysis; the occurrence of failure in this region as limited to the high hardness gradient
VALUE OF C, 0.090 0.160 0.216 0.314
transitional depths.
0.293 0.278 0.252 0.211 0.179 0.154 C.107 0.082
0.50 0.60 0.75 1,00 1.25 1.50 2.25 3.00
'I
4-4.2.1.2.3.2 Classic or Pitch Line Fatigue Classic or pitch-line pitting has been treated extensively in the literature and is related closely to classic bearing fatigue. Pitting life may be calculated as a function of Hertzian stress S,; it is a phenomenon associated with rolling contact, and the theory is not applicable if surface traction or shear stressws are of considerable magnitude. Consequent'y, valid aly arc limited to full EHD lubricant film
AMAXIMUM VALUE OF ORTHOGONAL SHEAR
STRESS OCCURS AT DEPTH - 0.33 b FOR CYLINDRICAL HERTZIAN CONTACT
41
83
r~l /Rc '--
---
2o
62D655 - R, 60 TO 0.020 in., 50 TO vUAin U3 in., CORE - R, 38 AAJS NITRIDED A&% 64.5 - Rc GDTO 0.007 l. L*s w t•S
-....
1
10
1th-41
640
50
OO
002
0.20
,5
0.20
0.040
000
0.Ot0
0.070
0.050
0.090
DEPTH BELOW SURFACE, in.
Figure 4-25. Case Depth Allowable vs Subsurface Shear 4-45
IL
-
1
-
The stress factor X is delined as
-I-''
15 LK I2~
-
-
,,..-
-
-
-
-4
mi
WF (-!
j').ps
(4-50)
where
total effective tooth load (Eq. 4-29), lb
Wf*; -
-d
pinion pitch diameter. in.
F
I
-face width, in. gear ratio N /N,. anumber m .-1.0L In mEq. 4-50 the term (M1,+ 1) isused with counter-
j~----*formal teeth and (min 1) is used with conformal 11 Ll1~-1teeth. --
-
-
-
Thec ffcctivc tooth load W,' varies from that
2S
39
15
40
455
used in Eq. 4-29 in that W, is taken at the pitch line. and for miost applications Wj - 0 because thc pitch load is governing while an incremental dynamic usuAlly is limited to the initial mesh contact. The stress index modified I in Eq. 4-48 is defined as
0
EQUIO'1NU tHAftQMC$$R,
SUL'L
4-6. vsHardessline hearYlel Figue F~gue hearV~el 4-6. vsHarm...load ftpaation as depicted in Regime 11 of Fig. 4-3. Opsrating in this region is not observed often in hedicopter drive subsystems where low-viscosity synthetic lubricants are used, except in the very high imei vA'~rino stna~es Tntm' veladitie&'V.. - V. + V. nf
the ordt.r of 15,0UU ft/nun or greater are rcquirod to achieve Regime 11 coaditions. The fundamental AGMA approach to the calculation of pitting life is given in Ref. 80. For helicoptcr use the following adaptation is suggested for calculation of the Hertz stress, S,: CP TLcalculated (4-48) psi
,
-
whom C. - elastic coefficient, (psi)i i - streas factor, psi I - stress index modifier, dimensionless In Eq. 4-48 the elastic coefficient C, is given by
4W --L
where k L
-
-
gometr ý.-Ylindrical contact k combined givent as E'
wherep, are the Poisson's ratios.
"6~
9)
jl~p
/
k
(4-)
where
cat.si
O'ess (4-52) 2 tases pftn rsueagc drasersgpaig psueage and the valegfrtecnatrti atrm loi us in par. 4-2.2.1.2.1. The value of S, calculrAted using Eq. 4-48 shuuld bc used with the S-N curve shown in Fig. 4-12 to predict pitting life. C1-
4-212.1.2.3.3 Wear Initiated Failure This is the most frequently encountered failure mode, predominating throughout the transitional lubrication states between pure: boundary layer and
factor, dimensionles (for contact k - 2.0; for elliptical 3.0) modulus of elasticity, psi
2 /+ '"+ E- +
' C~ t
full El-D conditions (Fig. 4-3). Causes and cor-Y
rective action are discussed in detail in pars. 4-1.2.11.2, 4-1.3.1, and 4-2.1.1. In the absence of relevant test data or cxtensivce xperience, the best procedure for analyzing this failure mode is to calculate S,.by Eq. 4-48 and to apply this value to the applicable S-N curve of Fig. 4-18. More suitable life equations that take into account many of the gignifierant variables other th'an ,~myso become available from the many research programs now under way. One such program, entitled "Relationship of Lubrication and Fatigue inCo~icentrated Contact", is being conducted by the Rescarch on Lubrication of the ASME.
.Committee
)
~AMCP 7W0-2W
4-2.2.13 Geer Drawing ail Specification Without a drawing or specification adequate to insure control of the critical variables, little confidence can be placed in the value of gear analyses relative to expected service performance; reliability goals cannot be guaranteed and the results of any specific Airworthiness Qualification Specification (AQS) test becomi relatively meaningless. To achieve a workable logistical, maintenance, safety of flight, and otherwise cost-effective helicopter program, consistency of product must become a paramount consideration. Consistency or reduction of variability is of far greater importance at the operational level than is the achievement of any other criterion of performance such as power-weight ratio, strength, or efficiency. The gear drawing must be amplified by numerous supporting specifications. However, the decision as to what class of data falls into each category is a matte" of individual preference provided the result is a workable system for procurement, quality control, and necessary engineering review and change. The 3 ,.hrZ;no ;a th. Aru-ntmi-nt
anvernino definition of the
component, and it must clarify any ambiguities in or between supporting documents. "Thefollowing review list is intended as a minimum guide for assuring completeness of data, but no stipulation is made whether it be provided by drawing or by specification: I. Raw material: a. Chemistry b. Certification condition c. Grain orientation d. Processing requirement .. si.c Shape and:
,
S'
rn.. . duction, fro"
in-os
f. Finish g. Decarb limits. 2. Heat treatment requirements: a. Process controls b. Certi!,ation c. Properties, including microstructure d. Case. hardness, surface and gradient, case depth r.nd tolerance, and core hardness, r. Quenching and tempering limitations including time, temperature, and interval regulations f. Limits on reprocessing, 3. Serialization: a. Proper identification and traceability b. Location of cod" and numbers c. System for transfer during processing d. Control of marking methods, size, and point(s) during processing for application. 4. Drawing technique: a. Specifications and Standards (MIL-D-I000),
MIL-STD-10 with dimensioning practices to ANSI Y 14.5 b. Gear reference axis definition with location tolerance for inspection set-up c. Specified taper, waviness, roundness, concentricity, and finish requirements, assuring compatibility for journals d. Boundaries to csed areas. 5. Finishing requiremeints: a. Specified methods and limitations on use b. Specified peening techniques including setup, shot, gaging, coverage, and certification frequoncy c. Means to avoid emlrittlcment and stress corrosion in all electrolytic, acid, or caustic proccsscs. 6. Stock removal: a. Limits on stock removal (minimum and maximum if required) during grinding on all cased areas within tolerances compatible with Item 2 and with design stress analysis b. Specified methods of control. 7. Nondestructive testing: a. Specified requirements and methods for
magnetic particle, pcnatrant, and cichant tcsia b. Specified frequency and sequence c. Specified frequency of certification of processes d. Specified equipment and precise location of identification for necessary hardness measurements on critical areas. 8. Balance require,.ents: a. Planes of measurement, limits, and speeds established and located, and permissible techniques specified when dynamic V'iiancing is required h.
Spr-ifdA
lw ation
limitc
and
material
removal methods for meeting ba'anct requirements. 9. Tooth form: a. Provide clear cnlargcd detail of tooth form, graphically specifying tooth thickness, flank and root finish, over pin (or ball) dimensions, OD root diameter, and minimurn fillet radius or equivalent b. Applicable data listed; i.e., N, Pd' "0' Dp. circular pitch ,P, involute base circle diameter DI, and %/. 10. Involute data: a. Slope and modification zones specified, e.g., by use of degrees rolled off base circle b. Critical diameters such as start of true involute, ard edge break limits defined. It. Lead data: a. If applicable, slope and crown defined b. End break limits and blend specified, 12. Allowable errors: a. Specified limits of manufacturing deviation 4-47
"
....
-.
from the desired b. As neceasary, equipment or certifiable equipment capabilities required to measure such
the primary failure rate of cylindrical roller bearings. In order to achieve the failure rate reductions required by modern MTBF goals, it is advisable that,
errors specified c. Repeatability and standardization of proof
as a minimum, the desig:,er: I. Come to agreement with the bearing supplier
check methods and frequency for inspection
with respect to specific application neetds.
equipment check specified d. Gear mounting within location limits given
2. Clearly specity appli-.ation requirements by pertinent drawing or specification,
in Item 4 or equivalent specified for inspection
3. Evaluate the effectiveness of potential gains
c. Tolerances on adjacent and accumulated tooth spacing, profile slope, lead slope, hollow or fuliness of profile and lead, waviness of profile and lead, and undercut o; cusp specified where appli-
available with amended specifications in order to undcrstand what changes in price are justifiable. 4. Inspect bearings for compliance with specification.
cable. 13. Chart format:
AMCP 706-201 describes the elements of bearing type selection and gives many examples of typical
a. Inspection chart for involute profile and lead required to conform to a predetermined standard for proper interpretation and consistency b. A sample chart with explanation of interpretive technique specifying magnification and paper travel speed provided, 14. Pattern limitations: a. For spiral bevel gears beating pattern checks required in lieu of profile and lekd checks Sb.Methods and ma'hine.s by which bearng or contact checks are performed on production corn-
helicopter configurations. The primary function of helicopter bearings is to provide accurate positioning of gear and shaft components under wide ranges of speed while also exhibiting satisfactory life. Means of achieving this goal are described in the paragraphs that follow.
ponents run against "working masters", checked in turn against "grand masters" specified). c. Data defining gaging dimensions, pattern size, shape, and location; and boundary tangencics
through specified V and H and profile settings specified (See Ref. 81 for further definitions). 4-2.2.2 Bearings The discussion of bearing application design, life arnaluysis, and draw•i1 ,otrois that fllows is limited to radial ba,, angular contact ball, and radial cylin d rical roller configurations. However, the basic principles introduced are sufficiently general to serve well in application design of any rolling element type bearins. Efficiency, reliability, survivability characteristics. and standards recommendations were treated prcviously in pars. 4-1.2.1 and 4-1.3.3. Army helicopter transmission bearings have exhibited a primary failure rate two times greater than that for gears and four tinme higher than that for all remaining transmission components (Ref. 46). Also, their replacement rate wt overhaul was three "times that of gears and 15 t~nies that of the remaining components. The majority of these replacements were due to secondary failure such as debris ingestion and corrosion. Also of importance to hhe designer is the finding that ball bearings (predominantly thrust applications) exhibited ten times 4-48
".- ZVI
4-2.1.2.1 Application Design Four aeneral areas apocar to create :he major difficuitics in bearing applicaiion tiasign. They arc.: I. Mounting practices 2. Lubrication techniques 3. Internal characteristics 4. Skidding control.
4-2 2.2.1.1 Mounting Practices In most helicopter applications of the rolling clement bearing, the loads are relatively large in relation to the physical dimensions and weight of the bearing. Good design requires consideration of the elastC buhaVioF of SUe, a sys"CM; adeQUate SUPPOr for both the rotating end nonrotating rings is necessary, and the supporting !-hafts and housings must have greater rigidity than the bearing rings. This criterion may be satisfied by use of shaft wall sections that are at least as large as the bearing inner ring thickness, and or total hotising-lincr-quill cross sections equal to the total bearing cross section. Use of thinner se..ions should be avoided unless careful stress and deflection analyses prove that they are feasible. Fretting wear, creep, and spinning are undcsirabie phenomena generally associated with the bearing inncr ring-shaft interface (inner ring rotating with respect to load vector). Proper inner ring interference fit is the most important parameter for control of these conditionFretting wear is the result of localized r.'bbing of very, small amplitude at the interface, and is difficult
(
*the
4There Zaused MR.described
)
to control unless the ring cross-sectional thickness is ring sections such as ure found in 200 or 300 series iargt enough for thc loading conditions involved. bmaings. TherF.ore. when dealing with appliction Of IkitAg croep is the x!aw rele~dve (lagging) motion of this formula to ligater section bearings combined ring with rtspect to the shaft and occurs in some with hollow shafts, it is nocuesacy to cnpfensate for irntances as a result of sufficieqi fretting wear to rethe reducmd radial pressure per unit interfceent* by duce the interference fit. A relative rotational speed the method of elastic i*ug theory as defined in Rcf. of as little as 10-6 X shaft speed may result in suf83. Use of interfercncc fits that produce surface tenficierl wear over several hundred hours to rediuce the silo stresstz in thc circumferenitial dirucion above 10,design interference. Often, creep o~cur-. initially 0CNQ psi should ibe ap-,%wched with -caution because bem~use of insufficient design interference, the fatigue life of the race may Ne reduced. Spinning is a term used to dencte an advanced 2. Eximssive thermal grsdicnts should be avoidc~d. statc of creen that occurs in cases w~h loose fir up or Becuse the circulated itib: ication oil acts to nodu1no interference between the shaf~t and dhe inrer ring late thesc gradicnits through forced convective of the bearing. With hardcne4, and groun~d precision cooling, an increase of oil flow to the shaft and interface surfaces, poliulhing and advanced wear rates beari-ng often can be used to alleviate thermul proboften result when the relative rotational spm 64aplems. proach 10 to 20% of siiaft speed under operating con3. Very thin ring sections es used in AFJPMA sizcs eitions not unlike thos'. in a simp'c slec'-e bearing, below series "0" for ball-type bearinas and series '"I Fig. 4-27 (from Ref. 82) shows a wear vs time funcfor cylindrical roller-type bearings should be avoidtion for a cylindrical roller bcaring application. ed for nigh-load applications. Elimminsion of the inare a number of design practices that may be ner ring by use or integral shaft raceways for cyto counteract these phenomena. They are lindrical roller applications is aii effective means of here in a descending orilew of preference, avoiding the problem) altogct~mcr. _rond. or hone' orn' tow) .-i umiardened 1. Thi exterit radhciattl~ i provide sufce rda t e pr[vent cree ma (roughness !5AA 10) with dimensional tolerance sufinLcforcetoI to creepimayterng orrun fIh baigwl be calculatedI and employed in the design. Two faceurtor le hnthrngofhebrigil tors that must be considered areassure attainmntca of dcsired c~alculated pressurets and a. Crcuferntil sreth ofthebeaing~'ig tili withstand frequent assembly and disassembly under applied rolling element loads which 'c~ctivcly adlogsrieitmnmllssfitreen. increases the inside diameter (ID) of the rind. 5. It is desirable to use 2/3-lip-depth shouldcrs, b. The influence of the temperature gradient spacers, and clamp nuts that are square with the jot-.rfrom shaft to the bearing ring upon relative iherwal nal surface end provide iigid axial clamping. Positive expansion. This gradient is a funct;-n of the cooling nut lock devices always should be used becAuse ring paths and of heat generation or friction loss. creep under high axial loading may rotete the nut. Ref. 6 Presents a calculation technioue suitableI for Deeiisse bearing ring creep results from a lag of the interforence fit determination. However, the ring behind tihe shaft ro, aioal speeoc, il is simple to calculations cited assume a soliJ shaft and generous dr~'erminc whether such cor~dition! will servc to tighten or loosen the nut. 6. Copper or silver plating on the shaft interfa~ce 4 ~ T as bAc used with wine beneficial results kp reC'~ 4'1I[A R~[~ ducing or virtually eliminatisig fretting corrosion and _____
.
-
thus prlnigcomponentsevcli,
CO ihI
SI-4-
~ -
.~
j
-~
-
I ---
I
.
4
---- -
A
4
-
-~
:0 ~
70
MI
yI
cvr between the shaft-key-bearing wing surfaces to
precipi-tateebe~nding fatigue failure inone or tacre of
~'~~-j C
-
7. Positive rwug-shaft interlocking with notched rings and keys have beer. employed to prevent creep. Howev;.r, satisfactory installations are difficult to achieve because statficient fretting corrosion may oc-
4
01'kATNCTIM. Nand
Figure 4-27. Creep Wear - Inner Ring Fit is Opersiliq Time
Outer ring (nonrotating loud) diametral clearances
clamping require somewhat I=m diligent at-
tention than dc the inner riiags of bearings. Howevei. heavil,' loaded angular contact thrust "~rings must be well clamped and their outside diameters
mutt be well supported to prevent excessive coning wumdr the action of tire aaglod rolling element toad victor, Whvri only radial loads or light thrust loads act involi~e6, retaining rings or similar devices arc vAtisfactory for axiail retention. The diamectral fit #eaerally should be nominally line-to-line to approximstely 0.OO~i in. tight at opetating temperature to reduce ring rottation. Outer ling inter ferencc !sm oftcn are limited by rtquirements for eawt of asserubly and disatisembiy. Higher speeds call for tfshtor fits. opostci~tdircottio Outr rng isnomaly tion to shaft rotation duc to thr rcinfrcs exerted by the loadecd rolling elements H owvrcein the~ case of lightly loadtd, outer-land-guided, cagetype bearings, the viscous drag may be sufracimnt to reverse the norm. t t oi ue citomry tscr.4ly ftte an Vined stee l'ilficx in aluminum or magnesium housings to reduce wear and the rate of iruireasc in outer ring snou~nun dear ance due to rising tempeiratwec The !ncreascd clearance atoeaiStmeauesol opertingtempratue shuld be coknpoclstod for wher~ the fit is specified at roomi ardMte ri Iraa t~~npe~~~aiorgll. he hec fit n amasivestel lnerhavrg he amecccficent of thermial expapaioin as the bearing will remain unchansed at operating temperature, but the Pt in a liaht alloy housing without liner will loosen in -proportion, with the product of bWaring outside diAmeter (0i)) temperature rise, and the difference in the thermal expansion cofrficlents of the alloy and Mcel, The change in the outer ring fit in the presence or a steel liner installed in a housing with considerabic diarriczae inlwrferen~e will lie between these tidn hnsindiarv
n,,ndiatin,,m snd
k cmealaidseO ea'silv On
Ulz assumption that liner is fitttd to a minimum of 250 dcg F intt~rfercnce (linc-to-livie contact when ýhc temperature diftfcrenial i% 2.50 deg F) at room temperature (par. 4-2.3.2), there is an appreciable, unifo~m prcssure at the iiner-housirig interface. This pressure resulti. in an e.lastic rvductioni in the diawrster of the. installed liner bore. As the temperature is inercased, Lac pressure is red-'aced and the bore cxpands. Ttic initial pressure, and, hence, the expansion rate, is dcpendent upon relative section thicknesses aund material properties that may be evalukated by application of the elastic cylinder whooey (Ref. 83). Fig. 4-28 preseasts a graphic sol ution to the room temperature fit correctian factor. To illustrate two practicabil extrcinet, coodi'ions are represenated both for a 140-mm-OD bWaring installed in a 0.045-in.-wall liner that is in turn fitted (250 dog F shrink- fit) into a 0.05in.-wall thickness alumiinurm housing; and for a 60min-OD bearing, with 0.090-in, wall liner (250 del; F 440
shrink fit) and 0.040-in.-v, H aluminum houiing. If the operating temperature is 220*F, the correctionr arc 000D13 in. for the 60-mm bearing and 0.0038 in. for the 140-mm bearing in this illustration. However, whe-n, as is often the case, the bearing is the principal heat soiurce and the housing provides appreciable he~et conduction, these values should be reduced to compensate for the temperature gradient. A rcascriaLle correction for most desigqn applications is 60%. Consequontly, the room temperature bearing outer .ring r'lt-rips should be tightened by approximately 0.0008 in. for the 60-mm OD bearing
I
baig dapoiaey002 n o h 4-wO Occasionally, due to space limitations or a desire to eliminate unnecessary detai! components, integral ex' ernal flanges are used on bearing outer rings for axial retention and prevention of rotation. In this case. care must be exercised to avoid radial restraint at the flange holes so as to preclude race distortioas due to thermal oi load-induced deflectiors. Under the influence of high radial loads, a bearing of this design always will exhibit greater stiffness at the ,,---. .1-~-as~~ nla --U~1
Ul
0.I R..UUMMS
M&Jr !a'
toward the flange. This characteristic can be used to homiensyatppfopriabedn delcation. of theig shift or thoseaing.b prpit oain fteiag ieo tebaig 4-2.2.2.1.2 LAbricaitoti Techniques While thioretically it is true that only a slight amount of oil is neceded to lubricate retainee rubbing surfaces and to supply fluid to rolling element-raceway conitinctions. helicoreter aninlicatiow, olýen 0"-3; 1,l quire the use of substantially greater -11iAti For example, circulating oii may be usted WOe zv retainer wear particles, water conden~aion uiid sludge, and to transport spalling failure debris to chip detectors or similar diagnostic aids. Increased oil flows also act to modulate temperature gradients by forced convective cooling and help to reduce ditlerenitial thermal expansion distortions that otherwist coald reduce component life. As speed.9 and loads increase, thermal stability can be attained only through the rapid rates of cooling provided by high oil flows. In addition, centrifugal accclerationG and windage barriers at heigh speeds make it very difficult to get oil to inncr raceway and cage lands, leading to a requirement for forced pressure lubrication, Finally, critical bearings often must be lubricated by redundant systems so as to increase operational reliability and permit safe operation should the primary system fail. rJenerally, lubricant applications may be grouped
-
0.607
S0.0D6
-1
140 mm00 BRG INBARE ALUMINUM HOUSING1 2 140 mm0D ORG. 0.045 in. LINER, 0.50 in. AL. WALL
3 60 mm 0Of. RG. BARE ALUMINUM HOUSING 4 60 MM OD BRG, 0.090 in. LINER, 0.40 in. AL. WALL -5 ANY 00 ORG. MASSIVE STEEL LINER1
-Ile' LU 0.004
*
--
TEMPERATURE
0.003
-0
I
-
TE0PEATUR
150
200
Figure 4-28. Tcn~wra'ure vs Outer Ring by approxim~att Dnz-values (diameter D, mim; mul.:I:-
A
-
__
_
e~ithei natural ot forced flow of a.Oil mist oil-laden atmosphere bWick feed to citber ring or retaniner c.Splash or dipping b) dammed oil level (ut lestt ceneer of lower folfirts Claumci) d.Gravity food &~tou~h dr310 or cust paisage from trap locatc4 to ca'sh ,iani, Saturn oil c. Su-facc tevuion aii/or conetirugal feed from rotating hollow shall with oi.l acquisition frot. pressure jtt or other me tti~c4R meann in consbination with premcdina atcýNtw f. Pressure jet stream inbpirgiur&on rcia'tir-. M~IS gap. 2. Moderate spezds (0.3 x 101 < Dr < 1.0 X l0fs): a. Lightly Ioadvr4 brai inj may rcspond well to 6z~ methocls of Item I
)
K)Q
"250
-Liner
Fit Redbction
b. Heavily loaded bearings require pressure jet -*ww.l
f ^
km sseh rinDho
a. Lightly loadtd bearings with relatively open faces can be lubricated by high-velocity jet implogcment b. Heavily loaded or restricted configurations rrquire internal pressure feed as described in Item 2. 0I1 ogre"s must be considered. Outer ring counterborod bali-type bearings and lipless outer rinb cylindricala frecluently are employed. 4-2.2.2.1.3 hitteuzIt Cbeaictevhstlc Of all of the internal geometric properties of hcariing the mrest impoi tank with rospedt to operating characteristics -I& dismetral clearance. Such factors as wnrarol of initial shaft displaceml-nt to reduce gear trisalignment, load sha.-ing of the rolling elosments in 4-31
.
*radial *duood *
*
j
load applicatians, elimipation of thermally inradial preload, reduction of externally induoed deformation load& (such as the pinch effect ofr ~pltnet idior Sears), and d'etermsination of ball bearing contact angle are basically depedent upon diametral clearance./ A specific operating diametriu clearance must be/ maintained under all conditions. While radial load/ ~deflection contributes to needed clearances, it is generally insufficitnt to compcrisate for bearing installatioci, or fit-up, practices or thermal expansion Chhdges in race diameter due to filt-up an:1 to
*
*
~~temperature differential between inner and outer
rings can be calculated directly from elastic cylinder theo'v as presented in Ref. 83. Whether or not an in.. crews in inner ring temperature will tend to reduce the raceway enlargement du. to initizl fit-up will idepend upon shaft temperature and heat flow con-/ ditions. It is nut uncommon to find highly loaded angular contact bearings operating at moderate to high speeds with the inner ring tcmi:erature 50-1004 deq F above the outer ring temnperature. Tolerances seieced for sh'tts hounig,qma ani -e-armigas w!il have a direct influ!:nce upon the success of cleat-ance compensation. Tht: range of variations of bearing de-flect'ons and lives in a given application and, therefore, the life scatter within a lot of ostensibly identical gearboxes, can be reduced greatly by use of bearings of the higher precision ABEC and RBEC classification. When high interference fit-up and clearance compensation arc required, there is a limiting practicable value for the ratio of ball diameter to radial cross-sectional thickness of the bearing. Ratios greater than 0.63 should be ap~proached with caution unless thcre isconsiderable experience from which to draw. The successful design of angular contact ball bearings for use in stacked seti of two or more requires a knowledge of their elastic behavior. DB (back-to-back; i.e., inner ring thrust faces opposed) and DF (face-to-face; i.e., inner ring thrust faces adjacent) configurations often are used to provide a combination of thrust and ratdial load capability, while DT (tandem) bearings generally are reserved for conditions where the thrust load is less than 4040%, of the radial load. Fig. 4-29 reprtsnts a single-row, angular contact bearing. Whtn operating speeds are such that the centrifugal force on the balls is noi sig;aificant and the ball-outer race loads are essentially equal to ballInner race loads, the line of contact is established by the centers of the race curvatuies. In Fig. 4-29 the radii of the races are denoted f, and I,(innmr and
4-52
.........
Li
// , -
4
I .
/
V21
'~ --
/ -
--
--------
--
Figure 4-29. Besaing Geomietry Change With lane( Ring Expansion outer, respectively). When a radial displacement of the iriner race A(D,/2) occur; due to fit-up or therm-Al growth , the radial clearanct: and contact angle .Are reduced. If the inner ring is allowed to displace -mially until ball contact exists with no load. a ring thrust face protrusion 6,. results. If a clariapedring DR ar OF mounting with a tight housing fit is employed, the resulting compression produces an internal preload and a comipansatin# inctease in contact angle. This may be eliminated by manufacturing each bearing with a thrust face intrusion equal to 6,.. The initial contact angle P. also should be reduced by approxiinazely (#,, - 0) If therr is a possibility for sizable thermal gradients and cesultant pieload, the DB mounting is preferred to the DF mounting because the load per unit of thermal expansion is considerably less with the DB mounting. Firt. 4-30 presents a graphic explanation of this condition. The thermal growth of x compensates frtaofYinBapicioswletervses true for OF miountings. The rcswaltant preloiad is an exporeaniai function of the relative Hertzian comnpression, 6N* Load-sharing equalization of DT installation may
DF MOUNTING
108 MOUNTING
Figure 4-30. Rtlative Thermal Preload
x
)
be enhanced significantly by requiring that the faces be flush (or equally offset) under a relatively henvy oxial gaging load. This load should be at least 25% of' the operating thrust load for maximum benefit. Raceway groove shoulder heights must be adequate to support the elliptical contact area of the 4,allrace. Shaft misalignment and combination ralialthrust loads can affect a skewed ball path that crntributes to this requirement. Most heavily loaded helicopter angi.ý.r contact bearings require a shouk :r height-to-ball diameiter ratio of 0.25. Excessively skewed patiss may require higher shoulders and increased radii of curvature for at IeaSi the nonrotating race groove. Such conditions also increase the requirem ent for retainer pocket-bali clearance to prevent excessive retainer wear or Irac.lure. The balls in the loaded zone of operation will pooition themselves as the race-ball traction conditions dictate, including highi meainer loads if clearances arc insufficient, Single-row angular contact bearings may be fabricated with two-piece inner rings (J-type bearing),
-
DF vs DS
and, in special instances, with two-piecie outer rings. This eliminates thg. need for counter-boring and permits the bearing to resist thrust in either direction. *Ihe inner groove generaly is ground with shims bet-aceon the split halves so that, upon removal of the shim and ausembly of the bearing, a Gothic arch shape results in the raceway groove. This configuration reduces the axial play of the bearing for a given contact angle. This fabrication technique generally enables one additional hall to be assembled into the complement, thus incremaing load-carrying capacity for the same exttunal envelope. Thrust-torudial lead ratios must be approximately 2.0 to p)revent degenerative three-point m!ritact. Misalignment and the resultant ball path skew also must be considered in avoid' ng three-point contact. The shim thickness that may be used in grinding the ring halves is limited by the need to maiintain race clearance in the absence of a thrust load. DWause the loaded inner ring is oaly half the width of that used in a conventional counte.bored ring bearing, total radial pressure between the loaded ring and the shalt 4-53
TA
2~ Is conadarably less for the Wa. intefrfeence fit-up. Conusquently, fretting and ring crewp also arc; inore difficlt to Control. 44.21.4CeruIcontact ~dd~ "I.M.4SWWO4 Ce"This Lightly loame hh-peed bearngs" may Operate withgros bewee sidi th rolin elinct 06Th pkcnsea and the rotating inner race. Such operation can produce snearing or race surface failures not unlike those camWe by gear tooth scuffing. The contrifusai aoaleration present at high spe crae a con sidmbe rllftelement/outer race load, with braking tr.iction forme exceeding driving traction forces at the rolling elemnent/inner rame contact. Retainer drag forcm and lubricant viscosity also play an important part in determining load-speed..lip con ditiolts. Historically. this distress mode has been a greater problemn with cylindrical roller bearings than with ball bearings. Calculation of slip-critical conditions is relatively uncertain, but some useful insight may be gained from Ref. S. One method for ;reventing gross slip is to maintam t te inerracerolingeleent ontct he oad re ,rn nlitsin surufieni ditrvine traction- 'Ms may be accompimstwaJ on DO5 or vi- angular contact blssnags with internal p.-elosid. A preload spring may wth bal anularconact ingl-ro be rquied tc ihsigerwbalagla bereqire
may require balancing to obtain satisfactory operalion. As in demernt skidding, a critical spoed exists; the centrifusal acceleration at this speed will displace an out-of balance retainer off center until all land occurs in a single local torse of the retainer. may, in turn, cause rapid retainer wear at the pockets is well as the guiding tails. Once started, the wwa rapidly accelerates until fallur occurs - often in 20 hr or tess. Hoeeadcotm fe xss ihrsett Hoceadhtmyftnxsswthrpcto clearance requirements. While controlled reduction of internal clearance to minimal values tends to redusce the skidding tendency of lightly loaded bearings, it comprises their ability to operate without lubrication, i.e., fail-safe operation (see par. "..3.Emergency Lubrication). Normal heat distribution within a bearing %rith inner ring iotation raiults in a negative temiperature gradient from inner ring through rolling ceincfltita and outer rnag to the housing with the inner ratzway Operating broadly from 50 -dkg to 100 deg F hotter than the outer raceway. Thc shaft and inner ring beat flow paths offer less rejection capability than the outer ring &adhousing paths. This, coupled with the customarily higher heat generation rate conefra w onit' neracts s~lidin th hi~v~her inne coteomacoatsrsusinheihrinr race operating tomperatures. Under normal operating conditions, the lubricant removes the bulk or the
Thene are several methods for contfolling gross slip wihclnrlAolrbaig wit Octlifdrical router acwaysma b eplye to produce a pinch effect upon installation. Installa4*io orientaiioa is required such that the external -War loads are orthqgonal to the pinch load plane (Rai. 84). ridiallv tiahtI. A small nurnkW af Mveiia* holwrollRm may be dispersed at even intervaLs througomt the coan~ipemet. This method requires areul uaimsiof the hoflow rollers to preclude bend*a faguen faiur ("e. 63). 3. Whim the eoafiguretion permits, the roller * * bearing may be mounted vory slightly off-center with respect to the shaft a16s. If the shaft is Positioned by an addiitional pai Of INearings, sufficient radial Pre*oad may be effected to provide the aceesary traction i1oad. 4. Out-of-round lerg have been employed to prodamc the same resul as in Item 1. The primary dia* advantaga to "hi tochnique is time inavased difficulty -in aseunbly and dlsaassombly. ITo miaminia skidding tendencies in high-speed tohkthe snallest acu4Aable diaimeters; should be -tWboth for the rolliftg element and for the pitch cir"do f hecompkmmeu of rollinS elements. Retainers
heat and maintains -thermal stabilization witbin this gradient. However, when the cooling effect and the friction reducing characteristics of the lubricant are absent. temperature stabilization can only occur at the higher gradients dictated by the increased friction and reduced heat rejection. If 'ntcrnal clearances are sufficient to accommodate the expansion attendant with the new gradient and increased cverall temperalure. then stable faii-safe operation is theoretically attainable. However, if inadtquate internal clearance exists, a radially tight condition results. This in turi! leads to a divergent increaise in temperature until bearing seizure or shaft failure occurs. As described previously, the mechanical means of providing positive rotation for the rolling ekLAnents in order to reduce skidding tendency can be applied in conjunction with greater iinertial clearance to affect a design without skidding and with fail-safe operating capability. Sinoe !he skidding tendency is highest in lightly loaded high speed bearings,. it is possible to install nonload-carrying hollow rollers in cylindrical roller bearings without loss of neoded capacity. This offers the dri-ing feature required to defeat siuioding "'nile providing adequate radial clearance to aacomnmodate thermal growth during fail-saft operation. Bearings with lower speed and higher loads cAlmibit
V
X
AMCP 706-202
progressively lea skidding tendency and are designed with adequate radial clearance for fail-safe operation without need for auxiliary positive driving features. Thermal growth due to fail-safe operation in angular contact duplex ball bearins can be accommoduted by providing adequate internal clearance initially (minimv'm contact anglt of say 30 deg) or, if initially preloaded, by mounting the bearings basikto-back (DB). back---back mounting allows the innet rings to gvow ra:ially and axially without generating additional prelad, i.e., radial growth tends to increase pveload while axial growth relaxes preload.
hence ignores any possible effects of gross elastic changes of shape in thew bodies. 3. Empirical coefficents used in the AFBMA formulas reflect the characteristics of air-melt AISI 52100 steel of Rc6O nominal hardness operating in medium-viscosity mineral oils at relatively low temperatures and moderate speeds and loads. 4. Any effects upon life caused by speed of rotation are omitted. S. The calculated lives are based upen the number of cyclic stressings to produce failure in 10% of the population of a statistically significant sample size. 6. Bearings manufactured by different sources arc assumed to belong to the same statistical population.
4.2-2.2 Life Anulysis Modern techniques for calculation of fatiguc life of bearings are based upon the pioneeri.i theoretical engineering and statistical analyses of Refs. P6 and 87. Certain empirical constants in these analyzes were determined by evaluation of experimental aata; hence, it may be argued that the effects of certain physical phenomena not specifically addressed in the
this is a valid argument, it follows that the bearing life in an application that dilers substantialiy from the laboratory conditions could vary significantly from the calculated value. Fortunately, the statistical model used (a modification of the function originally presented in Ref. 88) is sufficiently general to permit meaningful interpretation of faiiurc modes as diverse as human mortality. 1ght bulb filament burnout, or wear-initiated gear tooth spalling. Conscquently, valid test and field service experience can be used satirfactorilv to add life modification factors with corrected dispersions to the Weibull distribution for bearing lIfe prognosis. ".42.2,.7.1
S"•g
/
Asutptioos and Limitations
4-2.2.2.2.2
Modification Factor Approach to Life
Prediction A useful method h&s been advanced (Ref. 89) to account for many variables common in modern design applic.tions. An adjusted life LA is calculated as the product of adjustment, environmental and/or design factors, and the AFBMA calculated life Lio. LA - DEFGHLo, hr
(4-53)
where D - material factor (reflecting actual steel chemistry and purity), dimensio~nless E - processing factor (accounting for CEVM and other melting practices, thcrrnomechanical metal working, forging grain flow orientation, and absolute and clcment differential hardness), dimensionless AF lubrication factor (considerineg lubricant EHD film formation and relative surface
The basic AFBMA life calculations commonly used in the U.S. are based upon Refs. 86 and 87 and hence contain certain key assumptions and limitations: I. The failure mode is subsurface-initiated pitting or spalling. Cracks beqin at microscopic weak points, most probably at the depth of maximum subsurface orthogonal shear beneath the Hertzian contact. The developed solution, therefore, is based upon stressed volume theory. However, it has bw.n indicated (par. 4-2.1.1) that a preponderance of the failures in the analysis of helicopter bearings at oveih,- were stirface initiated,
roughnesses), dimensionless - speed effects (considering centrifugal acceleration and slip conditions), dimensionless H - misalignment factor (applicable to crowned and cylindrical roller bearings), di. mensionless It is not uncommon in helicopter bearing design for the value of the multiplicative group of factors to vary between 0.3 and 18 due to the range of conditions and requirements encountered. Digital cornputer programs often are used to define factors F. G. and H; while factors D and F are assigned values whether the life calculation is by simple AFBMA
coprcssive deformation of contacting bodies and
equation solution or by computer analysis.
2. Hertzialn stre•s theory is based upon :.he local
co~n",4-55
G
4-2±2..3 Compklee Elhadc sod Dynaimic: Solutione Dynamic forme associated with high-speed operation not only change bearing opra~ting characteristics greatly from those assumed for the static design, but also impose limiting speeds based upon failures due to sliding or con~tact slip heat generation. Fig. 4-31 shows that centrifugal acceleration at high speds not only increases the outer ring/ball *load for an angular contact bearing, but results in different contact angles at each race. The definitive axis * for ball rotation is dependent upon fthcontact thatt has the greater "grip" on the ball. At high speeds this may be the outer race, which then forces the inner race contact into gross sliding. Also, because the ball rotation is not coincident with the bearing axis of revolution, a gyroscopic procession moment is induced. For balls of large size and high contact angle, this moment may induce complete precession slip, often with immediate overheating failure. Analyses of the governing forces are treated in Ref. 90. General computer solutions employing the equations of this reference also may consider the elastic deformation 'of shaft and housings in combination with the Hertzian deflections between the race and the rolling element as they influence the load dstribution amnong a
number of individual bearings on a common shaft. A special case of elrztic deflection influence upon calculated life occurs in planetary idler bearings whose outer races are integral with the idler gears. As a result of the squeeze effect of the sun and ring gear radial load components, and of the moment upon the gear centroid d-ie to tangential tooth loads, considerable deformation occurs and may create additional bearing loads of sufficient magnitude to redine bearing life significantly. Rim section properties and internal clearances also have strong effects upon resultant life. Typical functions are shown in0 Fig. 4-32 (Ref. 91). 4-2.1.2.3 Drawing Contiols Confidence cannot be placed in the reliability or performance of a drive transmission bearing without a thorough evaluation of the important characteristics of the bearitig as defined for the specific application. Bearing characteristics may be controlled by drawing, secondary specification, or manufacturers' source documents - depending upon individual prefcrcnce. The following are minimal guidelines for such control: 1. !.1Mw
mincrial:V
a. Chemistry
b. Method of melt c. Certification limits d. Size reduction from ingot
OUTERRACE ELLIPSE
PGERTRTCONTACT LOAD ZOE
LOAD 0..~(N
f. Thermomnechanical processing limits if ap-
TOFO
P,. OUTER RACE *THRUTA
CINE GSRPIN)C
A
ia. OLN
roesscnrl bfPR . DCeabrtifation.
c. Properties rncluding microstructure, hardness d.CsP. crPrprt. hr apial c. L.imits on reprocessing
OWT
~--f.
-~
Retained austenitc, where applicable, or
--
G.
--
O
Dm
Ncq3.
Pa P1.
PURESLIDING
Serialization and identification: . Traceability b. Location of codes and numbers c. Process step for application d. Match marks for high pcints of eccentricity
on precision sets
IILLIPSIL CONTACT OWIRNRACE
e. CID code marking for verification of proper staczking of matched sets. __________4.
Figure 4-31. H~gh Speed Ampler Contact Dal TIMMuI SekariNg Forme 4-36
c. Grain orientation
SPINCOMPONENT
P" INNE ROACE
Dimensioning technique:
a. Applicable ABEC and RBEC grades b. Pitch diameter; rolling eleme-.it dimensions. race curvature ';contact angle (unmounted); radial clearance; shoulder heights; flushness and gsjing
(
F
z.
LUJ L6-
_
7j. C
RIGIDOUT1 R RACE
.
Cs
25I
0
0 .02
0,001
0.003
••
UM1
INTERNAL DIAMETRAL CLEARANCE, in... Fipft 4-32. Cleatace vs L,0 -- lIastic mA RilSk So~utdow
cluding pocket clearances, surface finishes; and apFlicable special dimenisions which differ from ABEC/RBEC standards. 5. FinishinS requittments" &.Methods and limitations on plating, peening. honing, polishing. and stock removal, when applaciable b. Protection aigainst erabrittlemrent and stress corroion., 6. Nondestructive testing: a. Requirements for maglnetic particle, pent:stint. and edchint tew~niques b. Control fivluency and sequence of test or inspection c. Fr-equency of cartirLiztion processes •) Tin as of fife modilrKation factors (Eq. 4-53) canunless specifically warranted or tubstentinted doumnt Inot be
-•
Basic introductory and clm~assfcton information concerning splin, his containedi in AMCP 706-201. Therefore, this discussion is limited to spocific desiln applications or power-transmittinll splints. The primary failure mods for a properly dnsinod sphn€ is wear. When trctvc motion is slight, Wmting corrosion often accelerates wear. Goo.d dkmn practice will insure at wear life in excess of the usdul comportent Mie. Galling and picku~p (weldin) occur only under excessive comprewive stress in the mecc or slight motion. Tooth breakna and betilum seldom ocin cur unieu shaft beandng momentsa•m aq-• the dtiign analysis. Nqql~ea of proper' fillet radius control or advanced frettinl corso oftenm contributes to iiuch failures, Blursting of the internally to_-:had member is rare, but molts from an exis insufba~kttp stiucture that tooth thin cessively tooth separating and centrifugal forces. 4-57
10,
•:
Two basic splint types are employed in drive system design: face splines and concentric splines. 4-.2,3.1 Face Splins Face splines are typically used to couple two shafts or a shaft and a Sear. Flat-faced, tapered, V, or square form teeth may be milled, shaper cut, or ground - dependin, upon material hardness and accuracy requirements. The fabrication of high-quality, interchangeable, concentric ;ou'ts with uniform tooth contact is relatively difficult & exl.p":vc; consequently, they ate seldom used in drive systems. The most common face spline is the CurvicO system (Ref. 92). Curvic* splines are easily fabricated with pressure angles between 10 deg and 30 deg, although the higher value is predominant. Most design deficiencies result from inadequate localization of tooth contact relative to the tooth ctnter or from inadequate clamping means to rcsist the tooth separating forces and the bending moments on th: joint, ORegistered Trademark - Gleason Works 4-2.2.3.2 Coaceutdlc or Longitudinal Splints
*
*-
Splines are clamped to prevent relative motion. Light interference fit dimensioning of major-diameter or side fit splines may be employed to assist in attaining secure clamping. When radial loads dominate the joint design, adjacent mating cylindrical bores and shoulders with light press fits frequently are used to complement the spline; successful application requires close control of concentricity between joint elements. Floating spline joints primarily are used to accommodate axial motion. Diametral looseness and backlash must be sufficient to provide clearance under operating conditions. The choiw between majordiameter and side fit control normally is predicated upon concentricity and balance requirements. Majordiameter control is preferable for precision applications where rotational speeds or alignment are critical. While side fit splints of 20 deg or greater pressure angles provide self-centering under torque loads, their looseness may permit excessive component imbalance and eccentric operation under noload, high-speed operation. Conventional involute splines may offer ap-
This type ofjoint has its a d-beating- surtfub essentially parallel to the rotational axes of the coupled components that comprise the external and internal
A safe design value axial force A1 for oil-lubricatcd spline may be taken as:
mating elements. This discussion is limited to the commonly used involute tooth form. although other
A/ - OA O/Dr lb
types are occasionally used. The involute spline may be manufactured by any involute gear production method, in addition to methods not well buited to full-depth, h'Sh-strength gear tooth forms. Depepding upon the limitations imposed by production volume and precision, involute splines may be produced by millins. shaDing. shear cutting, broaching, cylindrical thread rolling. rack cutting, shaving; by rolling, bobbing, and form or profile-generating grinding. Relatively trouble-free applications are limited to misalignments of 0.001 in./in. and are either clamped or floating. Operating misalignment of the axes of the mating parts of 0.25 deg or greater under load requires the use of flexible couplings (par. 4-3.2.1). It is very difficult to obtain satisfactory wear life with floating splines operating with misalignm.nt. 4-2.3.3 Properties of Splins Involute splines are designated as major-diameter or sidefit, depending upon the controlling dimensional features. Minor-diain.Ater-fit splines should be avoided in all but special applications (such as with a weaker internal member) due to the excessive stress concentration caused by the sharp tooth-root fillet radius on the cxttraid member. 4-53
where Q
(4-54)
- torque, lb-in. - pitch diameter, in. Reduction in slip for= may be achieved by use of special lubricants, friction-reducing tooth coatings, platings (silver, etc.) and treatments; with ball splines; Or by ,.h itroducti r .. .. (par. 4-3.2.1). Nylon (Ref. 93) and epoxy-bonded molybdenum disulfide coatings are often effective. Floating spline joints also are used to provide a slight accommodation for radial, axial, and angular misalignment. Under these operating conditiots. fretting and galling wear modes may prove troublesonic. Their occurrence is difficult to predict, and determination of secondary effects and solutions frequently must await design development testing. Depending upon the severity of the problem and the design restrictions, the following solutions have found widespread use individually or in combination: I. Increased hardness and accuracy (generally a matter of gear tooth grinding precision) 2. Shot peenring of onc or both members (relatively high intensities sind surfacc texture modification are desired) 3. Use of dissimilar materials, types of heat treatment, and hurdnesses Dp
r;
AMCP 70W-202 4. Crowning of the external member tooth flanks and major diameter 5. Increased oil flow or othek lubrication improvemeats. 4-2..34 Splme Streqi Analysh AMCP 706-201 gives allowable bearing pressurts St, for various classifications of involute splines. These values reflect app aximate current practice in the helicopter industry and are defined by SA•,=
,
-"
2Q/ (DF). psi
(4.55)
where D - pitch diameter, in. - face width. in, Q - torque. lb-in. for standard SAE or ANSI B5. i tooth proporti where tooth addenda are one-half those of AGMA standard 201.02 gears. However, these values do not repreaent true bxanring prssures because the accuracies, stiffneuses, and gecmetric proportions ofrtypicai spiines combine in rupe the truc cointact rea to less than the 100% tacitly assumed in Eq. 4-55. The accuracy of splines is determined by the tole'ances as specified and the method of inspc..ion employed rather than by the method of manufacture. Splines may be gaged (go-no-go systems), gaged and partially inspected analytically, or completely inpecaed analytically in the manner of gears. The following values of the fraction of theoretical contact sachieved with splines of the various classifications are realistic.
SC,,•Usifi"WAt Gaged USASI CI15 (commercial grade) measured Gag Iand C 1.5 B.19 measu0.45 ANSI ANSI B5.15-1950C.5 Analytically measured SAE C1 .3 (about 50% tolerance of B5.15 CI.5)
ficulty and coats. Reasonable proportions for spline fae width F and pitch diameter D, arc: 1.0 for torque-transmitting 0.4 < FID,, applications 0.8 < FID, < 2.0 for lowation and alignment applications. Lengthwise tooth load uniformity can be enhanced further by adjusting the shaft diameters and wall thicknesses to secure matching torsional deflections and by avoiding excessive radial stiffness at either end of the spline joint. Stress concentration must be avoided by specifying minimum fillet radius values, chamfeting or otherw*ise blending tooth ends into the shaft section, and achieving uniform loading. For most helicopter applications, spline fatigue endurance is not the limiting criterion because oscillatory loading due to shaft bending or torque fluctuation is avoided by proper design. If bending fatigue is a design consideration, recourse to use of the modifled Goodman diagram (par. 4-2.2.1.2) with appropriate stress concentration factors is required. Static stress analyses must demonstrate a positive margin of safety boh for limit torque compared toamaterial yield strength and for ultimate torque coarpated to material ultimate strength. Limit spline maximum drime system torque is defineda as 1.5 m torque as 1.5 limit continuous torque and ultimate l a 1 torque torque. The following spline static stresses should be calculated in addition to the bearing stress values SW.;: 1. Spline shear stress S;
'.-4M•"
-' 3Dý
Contact Fraction
1<.4
0.2• where -y < 0.7 0.75 < 'y •0.95
The variation of -y within each classification is dependent upon stiffness, proportions, and, in some instances, the ductility when the design load approaches limit shear strength. An external involute splint with face width F ,-
D.13 on a solid shaft will exhibit greater shear strength than the shaft if D. (outside diameter of the shaft) is only slightly smaller than the spline minor diameter D,,,. Therefore. excessive spline lengths can offer littlo 6cnefit while incrcasing manufacturing dif-
Q
spline torque, lb-in. number of spline teeth circular tooth thickness, in. T, fraction of theoretical contact, dimensionless 2. Torsional shear stress (external toothed mernbers) S, -
-
16QD., u(,/ - _ D,-)
p
(4-57)
where D, - inside diameter of shaft. in. Dw - minor diameter of spline, in.
"4-59
•
I. Burnting stre (Intenal toW.d member) Sj: 0
MIN.h4IDE* NICQSMM)IV [ASE Of SWMACWRIt
4
AB
~m4L79Xtan#.Dais
7.093 X 10-o1mC~ljrPat tF~~f1(~*7~J
A ADSOUFI[
(4-58.)
auk* diaineter of splmne. in 04- eotslde diameter of splkw tooth~ member D(i + 2 X(boc-u rim thickness)1 i.uI X roumlomalsped, rma p ria &-atitely. lb/in.1 01-
INOLUTE
T
O VN
K *VOLUTE~
IK
_____
i
imiwt and uhimale moargiis of safety must be calculatedl using appropriate torque values in these-
equations and comparing the stresses calculated with Eqs. 4-56 and 4-57 with the allowabke yield and ultimate shear stram= of the respective parts. The sitreses cailculated by EF4.4-58 are compa"e with the allowable yield and ultimate tensile reususeof the intunaz sothed Memoer. *
4-1.2J.5 Drawling Desigum a C.arol Splines should be specified on the engineering *drawing or other document in a manner similar to gear teoth. An enlarged, dimetsioned sectional view and a data block should be included. An even number of teeth is preferred for over/under wire inspection purposes and manufacturing ease. To prevent .involute undercutting and permit use of the maximum number of manufacturing technique op. tions, tooth numbers must be no lower than those in Fig. 4-33. *The enlarged spline drawing (D,,/2 is aconvenient scale) should present major diameter, pitch diameter, form diameter, minor diameter, minimum M1llt radius, circular tooth thickness (external), circular space width (internal), tooth tip chamifer, dimension over (external) or under (internal) gaec wires, and surfacte finish, The data block should present number r-f teeth, diamectral pitch fraction a/bl(whcrt a repiesents P, D,/N and b is the value of P, when it isexpressed as the reciprocal of the addendum length), pzressure angle, base diameter, total or composite index error, maximum deviation of parallelism tooth-to-tooth for given length orfte3agemtnt. and parallelism limits with respect to part reference axis or surfaces. Involute proft3c !qlefianc2 should be Dreigsen!e4 data, for gage inspcvtion. or as a chart, for analytical Inipection. For sag*e inspection techniques, major diamieter and pitch diameter eccentricities must be
4.6
_
_-____
o____ X
?h%
PRESSR
NL,,4
is
VALUES NUMBERED) (REC~OWENDE
ilgure
4-33. luvomuse Swie
Bl. aai 0wVi
absorbed within the limitations of the mejor diameter and effective tooth (or space) thickness tolerances. For externally gagr4 splines the maximum effective and minimum actual circular tooth thicknesses must be specified; while for internally gaged splines, the minimum effective and maximum actual circular space width must be determined. When Inspection with gages is specified. the diameters of over-and-under wires called for are referenced oata. With analytical inspection techniques, tooth thickness and space width are given as actual minimum and maximum limits. and toothto-tooth spa.ing tolerances also must be specified. In addition, allowable tooth lead error should be sub,stituted for parallelism error. The manufacturing method must be considered when detailing the spline. Shaper cut splines should have aminimum chip and cutter overrun gap equal to their total depth. The minimum relief diameter for an internal spline should be equal to the major diameter plus one quarter of the whole depth, and for an' txtornal spline. the minor diameter minus one quarter or the whole depth. These diamectral clearancc values also may be used for broached splines. Hobbed or ground splines, of course, must provide overrun c12,siace for the wheel or hob radius. 4-L.24 Overrniniing Chicte. Certain overrunning (free-wheeling) clutch requirenients were described in par. 4-1.2.2. The lowest drive
-
il
"T.V"I.
L
• ••:•
N
a
AW-•P 706-202
system weight will result from placing the clutch in the location of highest speed; i.e., between the engine and the first stage of reduction gearing. HoNever. other considerations in multiengine configurakions such as vulnerability, safety, and reliability - may require locating the clutch between the first and second stage of reduction gearing, Operational requircments for clutches vary with helicopter configuration, mission, and life cycle. However, current Army clutch requirements for twin-engine helicopters typically art.. I. The minimum ultimate torque capacity of the "dutchshaIl be 2.0 x limit torque (limit torque - 1.5 X maximum continuous drive torque). 2. 2-hr, full-speed continuous overrunning shall be posbible without operational capability impairment. 3. 30-min, full-spocd continuous safe operation shall be possible after total loss of hlbricant system. 7hse requirements apply to the entire. clutch system, including support bearings and, often, seals. In addition to these requirements, there exist other significant design considerations as evidenced by observed failure modes in existing helicopter clutch ap-
Splihiiui,;
)
T.
I. Brineiling due to the presence of oscillatory torque pulsations. Depending upon operating stresses and configuration, overrunning clutches wilt safely tolerate only 10 to 30% continuous oscillatory torque. External shaft bending and radial or moment loads must be eliminated from the clutches by use of relatively rigid support bearings that maintain concentricity at all times between the driving and overrunning members. When modeling a drive system for torsional analysis, it is important to consider the clutch as a rclatively soft torsional spring. Stiffness values typically range frog 35,000 to 350,000 in.lb/rad (Ref. 94). 2. Excessive wear at intermediate overrunning speeds. Maximum wear conditions usually are encountered when the output member is operating at full speed and the engine is at idle speed. Most spragand roller-type clutches evidence their greatest wear rates wlhen (input speed) / (output speed) - 0.5 due to the product of centrifugally induced compressive stresses and sliding velocities. 3. Failure to engage at high speeds. In many instances reported, the second engine has failed to engage after the first has accelerated the system to ground idle speed. Both sprag and roller clutches rtquire a critical friction coefficient of about 0.05-0.07 to engage. Hydrodynamic or elastohydrodynamic oil film formation and/or externally induced vibratory modes may lower friction coefficients below this level at the moment of speed synchronization, resulting in
momentary or complete overspeods. Subsquwnt adjustments of input or output speeds may lead to abrupt engagement with attendant shock loads sufficient to fail adjacent drive system components. 4. False brinelling of clutch elements or support bearings. Clutch support bearings operate in a static mode whenever the dlutch is engaged because both inner and outer bearing rings rotate in unison. Extcrnal vibration thus may cause fretting or false brinelling at the roiling element/racc contacts. The entrapment of wear particles and sludge in the outer race often accelerates such wear. Therefore, it is important to maximize the static capacity of the support bearings for the available envelope and to provide good oil circulation without stagnation areas. Design considerations peculiar to particular types of clutches am given in the paragraphs that follow. 4.2.L4.1 Spaig Cts s Sprag clutches am the most widely used type for helicopter drive systems. Two variations have been used with success. Both employ a complement of equally spaced, full-phasing sprag cams operating between concentric circular races. A detailed study of their geometric and operating characteristics is presented in Ref. 95. Race cross sections must be sufficiently large to prevent elastic deflection under load from increasing the sprag space by more than about 0.002 in. Race hardness and case depth must be adequate to support operating Heriuian stresses of 450,000 to 500,0W psi at the sprag/inner race contact. Successful applications of these designs are based upon between 3 X 10' and 10' cycles of full torque application without failure. For ,noderate to high-speed operation it is preferable to use outer race power input with inner race overrunning to reduc the centrifugally induced sprag/racA contact stress. In such usage the sprag complement should remain stationary with respect to the outer race and should slip at the inner race during overrunning. This arrangement also permits centritugal-feed lubrication through the inner race and reduces the race/sprag sliding velocity for a given overrunning speed. Both clutch types usually employ a degree of centrifugal self-energimation by virtue of sprag center of gravity (CG) offset with respect to their ,.ontact engagement axes. This can cause some problems with high-speed applications because the drag torque (power loss) and wear may be excessive. The two dutch types differ in some characteristics. For example, one user two concentric cage elements to separate the sprags, while the other uses a single outer cage. The double-cage type uses an inner race 4-61
K
EP-705-2D2 drag spring to react the centrifugal self-energization
4-2.2.4.3 Self-energlzing Spring Clutches
during slip conditions, and may producv con-
Although there have been kio applications of spring
siderably lower overall drag forces at intermediate slip conditions. The single-cage design frequently employs an integral rib on the sprag which contacts the adjacent sprag ard thus limits sprag overload rock angle. The overload failure modes of the two types also differ. The double-cage design fails by sprag turnover (with resultant pernmanent loss of drive) while the design with ribbed sprag and a single cage fails by slipping. However, both failure modes gen-rally exceed the 2 X limit torque requirement by cornfortable margins. The single-cage clutch usually has a higher torsional spring rate than the double (for equal envelopes) along with a greater tolerance for oscillatory loading conditions.
clutches in production helicopter systems, considerable interest has developed in them because they have the potential advantage of reduced weight and size for a Liven torque capacity. The principal detorrent to the use of spring clutches in helicopters has been their poor release characteristics in overrunning. Recent design improvements feature a tapered-width helical spring of rectangular cross section and cylindrical outside diameter (Ref. 97). The torque transmission is between a cylindrical outer race and the outside diemeter of the spring. The device may be servo-actuated with an energizing pawl that contacts the small end of the spring or selfenergized by friction forces between the spring end and the outer race. Recent development and test cx-
Proper lubrication of either type requircs coinplete oil immersion. This often is accomplished by use of full-depth cirrular dams on both sides of the s;rag unit.
perience is reported irn Ref. 98.
E.L2.4,Z Ra-p and Roller Cl Ramp and roller clutches also have found cxtensive successful application in existing helicopters. Such designs employ a cylindrical outer race as in sprag clutches, but use cylindrical (hollow or solid) rollers in lieu of spra•s, plus a multiple-cam-surface injier race to provide a wedging action on the rollrrs upon engagement. O,,errunning usually produces roller complement roli.; contact with the outer race and sliding with the inmu... Consequently, most moderate- to high-speed applications feature inner r*e.re input with outer r~acc overrunning. This desion may require forced feed (pressure) lubrication through the inner race to obtain satisfactory fullsneed overrunning, Spring-loaded cages or individual roller springs are used to force the rollers into the wedge to secure reliable and rapid engagement upon race speed synchronization. A thorough analysis of the dtsign geometry and speed characteristics of these clutches is givea in Ref. 96. Due to the rcduced radius of curvature in the roller as compared to the spra8 cam, roller clutches have lower torque capacities than sprag clutches of comparable size. Ovcrrunning drag torque also is gteater at high speed& for the roller clutch. Failure mode of this clutch type at oveytorque is slip, if the cam/iuce components arc sufficiently strong to preclude their fracture, drive capability is not lost. In most installations, the roller clutch has shown a superior tolerance to oscillatory torquc-induced w-ar. 4-62
4-2.2.5 Rotor Brakes AMCP 706-201 describes the basic requirements for rotor brakes, while AMCP 706-203 presents the minimum qualification test requirements. Thi discussion, therefore, is confined to typical detail requirements and limitations and to basic design and analysis procedures. While this paragraph treats only hydraulically actuated disk- and puck-type brakes, the basic analytical techniques presented are sufficiently general to aid in the development of design criteria for other types of rotor brakes. The disk brake has become virtually the standard for helicopters due to its relahve simplicity, ease of inspcction and maintenance, and reliability.
A rotor brake differs significantly from a wheel brake both in failure modes and in functional requirements. The catastrophic failure mode for a wheel brake is failure to engage, or failure to stop the aircraft. Puck clearances are nil, contact speeds are moderate, significant cooling may occur during and after use with disk ventilation assisted by rotation, and repetitive use with short operating cycles and interva!a is common. The catastrophic failure mode for a rotor brake, on the other hand, is unintentional operation. Puck clca-ances must be very large, contact spreds may be very high, the primary cooling is provided by the disk heat sink, and iepetitive use in less than a 5-mmn time interval is virtually impossible. 4-2.2.311 Requirements and Limitations Recent performance specifications for Army heli-
'
alaA
706-202-.
-
reureets.
~
1.~~~~
wiehlcopter ro
~
Shl
rks i hnvt in uclue
n3 tprtrfom10 v
h
romiessu-t
cc-rtr(nua)
olwn
~cain
~~x
h
4. SMllsthopd rotor ftopped whil secked n3 are at
~
4-t. win 3. Must hndotb oratedoppn agmain rvst
W'I
tI(Jglr
eclrtordsc t lb-lcptrweg
rae (4-6(1on)ah
wheei, lb ei'foctivc coefficient of friction, gear to ground, dimension' As 6. Activation and control shall be fail-safe. - track of braked wheels, ft d Safeguards are required to prevent inadvertent RcliFor a stop time of 15 sec, rotor inertia I - 7500 sthigvation. Engine control interlock required with ft' and 0 -250 rpm rotor speed, Q5 13,090 lb fit positive raefeidon in lock and o:nlock modes. (ignoring aerodynamic rotor decay). If each of the of Deveopmnt ad prforanc hisofi braked wheels is loaded to 4000 lb, the wheel-track is and~ pherformawnge guidelines:o h eeloptentls helcopersals suges th folowng uidlins. 0 in., and f 0.4 (rubber sliding cm asphalt), the 1. The best (in simplicity, reliability. and safety) skid torque Q5 - 10,667 lb-ft. and therefore. a hydraulic system is a manual hydrostatic type. If dangerous ground loop potcrntial would exist. The operated from or boosted by pump/accumulator minimum stopping time tandet such conditions would systems, these -systems should be divorced comnbe 18.4 sec. A safe limit for tlhe pilot-activated rotor plptely Affoin 0fligh coigtroli or sur o actuator systtr~s. brake, rnudle would apply aboui 9800 lbb-ft%of torquc 2. Autorn!2ti ielf-vidjustment is undesirable beto the main rotor mast (equivalent to a 20-sec stop). cause it comprontis%~ reliability. Suficient fluid 2. Although the static brakaway ffiction for the should be provided to accommodate the uxsful wear dibk/puck brake may be somewhat higher than that life of the lin.ngs; martial hydrostatic, dual-level, for dynamic conditions, the severe coniequences of mechanical advantage systems have 4e.n developed inadvartcnit rotor rotation during engine idle operato accomplish this requittunent (Ref. 99). tion suggest the need for an additional safety mar3. The dick should be stiMy coupled. A short, oto- gin. This ean be provided by use of the lower value o' s'ionally stiff takeoff dave on the main rutor transfriction cocfficient in design calculations. Thus, a mission is often desira b~t. Soft mounted disks (such typical pair cf engines might develop the equivalent as on intermediate or tail rotor gearbox drive-shaft main rotor tvrque of 12,000 lb-ft at W0% gas hmngwers) invariably bneorne a vibration and antinode Itnerator sneeda. If thc safe vilot-activated rotor at borne speed duramg engagement with resultant osstopping move is limited to a w~ain rotor torque of 61llatory lands on disk and/or puck attachments. 9800 lb-ft in accordance with Item 1, a secnd brake Thewc loads may causf. intermittent brake chattc~r. y modc with increased pressure (intcrlock-protected for nanmic system ovvrloads, or crew annoyance, engine start soquence only) would be indicated. The pack or caliper asstmbly should have a high spaing rate lflouotiiig that it, stiff in all loading -ipitrtdretos vector comoetdietos 4-21.S Deelg s Analysis From. ii practical viewpoint, t~herrc may be iimiTwo basic determninations arc required for the caltation3 that place two or more of these requirements culation of safe brake performence; (1) limit energy into conflict. Often, specification compromise or rate per unit ares to yidd satisftictory wear Wie and multimode brake activation systems arc thi result. piu~udc disk scuffling, and (2) disk heat 6rnk capaFor example: city. Surface enery rate varies with such lining and disk 1. Short stop time, high rotor inertia, and landing properties as thermal conductvity, diffusivity, congear skid friction limits may combine to cause ground vect-vc cooling, and critical tempeirature. Solid stel S loop. disks in helicopter applications have been operated Braking torque Qp is successrully at an energy disaipation rate ED of 25 N Dtu/in.1-min. The referenced area isthe swept area M. Ilb-ft (4-19) under the I'uck. The wiergy to be dissipatedc is the where lining debris oould cause FOD to etgines; or
*APU.
f
-
.,
hk-.-4.
A
4-63
AMCP 706k202 kinetic encrgy of the rotor at time of brake application less any applicable rotor aerodynamic decay increment. Wcar life for common brake puck materials is dcpendent upon surface temperature, pressure, and velocity. Existing rubber-asbestos lining materials have dcaienstrated wear rates or approximately 0.0004 in.'/sEop at pressures of' 240) lb/in?1 for mean rubbing velocities of 6000 fpm over 20-sec stop pheriosecn- eemnto nvl h etsn Theecoddeermnaton ivoles he hat ink capacity of the disk. For stops on the order of 20 to 30 sec, steel disk thicknesses in excess of 0.5 in. offer little hc'p in reducin~g peak surface temperatures at the end of a stop auc to thc limited thermal conductui~y of steel. Current systems operate wel with vaiues of energy/pound- or-disk near .100,000 ft-ho/lb with peak disk rim temperatures of about 500*-(OO F. Other mnaterials such as beryllium and carbon graphite recently have been emiployed with relative success for hralze ap~plications. The greatest improv nient seems to be available with a configuration -.
which uses a piwprictary low modulus structural graphitc composition for both the disc and puck 'istu I niiai A beylu ciS Ici s uteu between 4ie graphite lining and the hydraulic siave cylinders. The saddle or caliper assembly ;. fabribctikesensu aa tSucbaed wihav d ro lm ndm ctedsfuly onstructed Such brakedihdsc thavben suto 1/5 in., puck diameters of 5 in., and disc diameters of 18 in. Considerable increase in energy storage, allowable operating temperature, and wear life has beern demnonstrated. Safe disc temperatures of 3000F (incandeicent white light) and a thermal capacity of 300,000 ft-lb/lb of graphite Eiv. the total stopping energy is dissipated in abhnivc wear. The puck material also serves as a ht at sir~k and maye in cnsiere he heral csig~ cpacty. The ildvantages, due to these characteristics seem to indicate that weight savings on the order of 50% and wear life increa-es of S00% relafive: to conventional steel/ rubber-asbestos sysiems arc obtainable. These factors are probably sufficicat to offset the initiak tuigh cost to the extent that a lift cycle cost reduction can be achieved. Disadvantages lie in initial costs and structural limitations of the graphite material. Although ballistic: impact characieristics, are satisfactory and handling damagz susceptibility is relatively low, graphite cavnot compare with steel. Through bolt or spline alttchments cannot be used for the graphitr disc - a high pressure squeez plate or friction drive attachment i,3 required. Similaly, thc disc cannot be. 4-64
drilled locally to achieve dynamic~ balance requiremerits. However, a steel reinforcing ring may be used for this purpose. The structural graphite material costs about 5400/lb as of this writing (1975) - future costs may be significantly less with adequate production volume. 42-23 STATIC COMPONENTS Static or nonrotating components of the transmission and drive system include the gearbox hosnliequlmutstdadows touenlos,andoupport sthes adymi owelservers thatins pnns.hipagrhadessolytemt signilicanThcomporents;hiae.,ecaes and y thousins, t adhuig, d sinfctcopet;ie.ca quills. 4-2.3.1 Cases sand Housings lillicoel~r gearbox cases mid housing& are fabriczied alit~est exclusive.ly from lightweight aluminmi alloys and castings and forgings or from magnesiun aoc~ns heemtrasehbi xeln hr Ths ty andteimalsehbte stexgeh-eo-tihtr ~ting v salond ratios, are readily machinable, and in many in-. stance may be salvaged by weldingi and stress reilieving with little ttmuli~ant iouss u *sircrigib propel te.Atog h eea pra~ ocsigad forging designk is well covered in avaiiable literature, pic-n r sm~ie nte aar sta certain aspects peculiar to Army helicopter apfolw 4-23.1.1 wi-i A~st6 Hcult~ c eai~ mayi be claissified as primary sl f tct uxa] lund paths (rotor masi supiort or control sysu;nt reaction member) or s~mpiy as Swir housings cant. This distinction is fundamenital in the selection oftediganaalssmto epyd.Cicality classifications of castings arnd fortings are. defic in MIL-C-6021 and are interpreted in AMICP 706-203. In most instances specified crash load fawors and limit maneuver loads will require ultimate and yield strength levels in primary structural cases and housings of suach magnitude as to permit design deft nition by static analysis as opposed to fatigue analysis. Fatigut analysis will be used to define only the rotor control reaction portions of the cases anid, occasiory.4y the gearbox support or mounting lugs when rotor vibrartory loads or dynamic reaction loads frem ground resonance or lardinig conditions are sufficicat to cause concern for low-cycle fatigue. SLa~ic and fatigue test requiremnents are outlined in AMCP 706-203, which descibes basic design load
_
AMcP 7n5-202 requirements or MIL-S-869g. Tine, critical design criteiton genecrally is the astisfactic-ai of the static test requirements. Because thc intckgrity of a casting or fni-ging is 8ovcrned by the type of quality control cstablished by applicable d.-awings and snecifications, it ib imperative that required static tests be performed on the [eut acceptable specimens. The radiographic acceptance stancard ASTM E-155. as well is other inspection criteria, ther may be based upon thecso static test results. Recent Army helicopttr RF1P requiremcuts have emphasized inc, ~asrd crew safety through more crashwortby de.,.,gn in accord with t1k recoinmendeiions of Ref. 100. Limit load conditions art based upon +3.5 aud -0.5 maneuver load factors at the helicopter CG and ultimate losad conditions upon normal load fartors of +20/- 10. and lateral or longitudinal load factors of *2(.. Combination loading also must be considered as the simultaneous occurrence of loadings in accordance with any of the three conditions that follow: Condition I
between all mectals except those immediately adjacent in the activity series (MIL-STD4S9). Excesuive steady tensile stremies due to assembly clamp ing (such as use of bolted devis lugs without spacers) shou!d be avoided to reduce the susceptibility to stress corrosion. 2. Lack or attention to differential thermal expansion. Steel bearing clamp nuts and similar devices installed in magnesium or aluminum threaded bores often lose thcir entire axial clamping force at operating temperatures. Such applications eithermust have an initial deflection that isgrcater than the amount of thermal relaxation or else threaded steel liners must be inserted in the case bores. Static bearing and hoop stresses should be checked throughout the possiblt ambient temperature range (no-rmally --650 to +3i00F) when steel and light alloy cases are joined with piloted flanges. Them ~ally fit steel liners in alloy %ascsshould have a nominal 300' F interference and the bore of the light alloy ring section surrounding the liner also should show a positive margin on limit stress at -65*F. Where steel bhearingst are installed in the liner, their line-to-line fit-
* 10 + 10 +20 _k 0 10 _k 10 Lateral Where a structural support casn isof relatively simple configuration, forgint.i arc pscferred to castings
up temperature and outer ring cross section must be sion in the housing bore fiber at -6ff F ass, ining the bearing fit is line-to-line at a temperature of 72*F may be taken as:
.~Longitudinal
because of the superior strength-to-weight ratio and
the inherently lower variability in strength of the
s_
a,)_W_+___)_36
)(2 (d' -e)
=
rormer.
'
The four dcsign deficiencies found most frequently in current %rmy helicopter housing com-
-
)
faces by thaiough cleaninig prior to ru~in impmg-
nation. \ Galvanic corrosion protet infel omk rsn \ arlic, or zinc chromate .i.--u.*,o should bL- used
(D Of
I +--61
tihle pcdn member beig larg rosion ousiing upper H + L±e dstdigns waster Suhvoidtra (ormay swleaterfo an, ashy wh re with reastude cas byl nounnabrcrdeig wal enc-iormia baigctrrn
compounds. Sharp edges and 6%w~h sbr (aces must be eliminated by chamfering, polming. or tumble (slurry) dc-burring to avoid ipadequatc resin or ps.nt coverage due to surface tesi ieects. Cathodic particdes must be removeJ cosnplci%.y from casting Sur-
I
IE
ponents are: 1. Insufficient attention to corrosion protection. life kL Success in attainment of senu;--_ compFnentw
6
X
-
b e d E a
oe n
bearing 01), ir. liner 01) (ateel). in. -light alley i.eetion 0M) in. - Young's modulus, psi - linear coefficient of thermal expansion, -
in./in.-*F it Poisson's ratio, diumesionlessi Subscript I - steel properties Subscript 2 - light alloy properties
"4-5
>.
3. Improper attention to joint and fastener toq"ireinuceis. Sufficmet flange thickness must be proevidto dtribute loads uniformly among the proloaded tension fastenr (bolts or studs) susJ on case flange joints. Fastas pmload must fi sullicient to "maintain tension at -651F. preclude strs mvmsls during normal ocaillatory loading, and maintain Vflane contact under tension loading. For ,opeuly designed fangs with compatible fat.cr spacing, a conservative value for fstener tension loading P, for monact-loadod cylindrical joints is given by P
-
M
(4-,)
b
where N - moment. in.-lb Doc - bolt cirl diamneter, in. N - number of bolts or studs (equal ipacing assumed) The yied and ultimate streagths given in MILIGIDIl-5 for tandard AN studs repesnt th , r ab•m ,htead wih rn-r= iAIh.___. -. on have a ofte Sowemcharactenst, more daletios efect upon the instaled arengtsof eaidl. AN aeude than upon lgW sm. Eves with caxius practicable pu Vpendnlaaiy for tapped
boles the combined effects of aper snd squarenew of joints may indwa benteg koad such that 3/&-im. srom studs demoasrate ueauflc filure at 90% of hd"ook minwimum values. The significance of th is probbly bete demonstrated by the fact t"at typil 7/164is. dimeu stdsusumlly demonstrat double the iintu tenile strength of comparable "t
*
rlrm. 1
":
.
-&.-A-
a6 WIM
-:4.
WE&" mitaf.
-.
•,si.'
-4
,
weight pnatie, of 10, and 25%. rspectively. S fltid tapsled holes for s• istalatioes m'um be vented. Whe toerqu must be transferred throuh a flang joins a ouqu capacity Q(4in the abeewe of externlr Unas loads is given by
*whew N bloc ,o iJn d Daking 4-"
a toqpueL bia. - ma4k of smu -boftckk disisrnwii. - stud •"diameer, in. this valet doeis, key or othir mcheacal devses, dsMuld be se to prevm beain
Although it is comand shr di of the studs. mon to consider stud sh stren4th in determininug ulimate joint streagth, i Sctor should not be Pie upon for normal design torque asnalysus The jo*L oaiguulon dtortically neessary to mt the strength requirements having been detmined, it becomes imperative to asure cornpatible dfta desig of the machined surane for the ftenrs. Fractur resiuta•ce as well as fatigue condlatkion requise careful attemtion to semirgly les sigicant deA. Onerus- filet radii must be providMed in spotfaces. ountnborns. and keyways. All rath edge -arp curre mud be ciamf•ered, a"l b must be removed to minim= mes ctcnnratsoa. Smootb blending of inateusctions on critical machined surfaues is also w iesay to mimimize stress coacesaratio. Addit"ony the ow rbone or sotfae eim must be adequate to provide wre.b clearsa. for normal mauatmacace opavion. 4. Failure to covmmt dsllectim vd loading in coftgous structure. Many re d istancu s of gerbox mounting lug failure am attributiale to x.ternally induace So*& that wer Miored in the design anhiv. The cmst &wbox structura whi ", often ii a suTir than she a-raum summon to wic itandis mounted may provide a load path for bending fow torsin ractions p -sent in the airframe strsaure due to boding gar, roto thrust, or vibratory resposess. Examples are the aauchm•nt of a four lug amrmy earbox to a dec strocture that has vatscal beading odes, or the smilar mount* of sa iniuo0du r genbos on tail boom sructure that usr tail rotor dallos torsional ddeiodas with to obtain the safpith inputs. Beam it is doesk ty inhert with mountiag raedndancy, it is usually a m A *%r am mfa,!is sneak.e wte mom rr - r' rP -C---,-frme strucue locally or t' Wirotta aduistaol compliamnc at out or mor,. e•arbox attachment by oanof elaomeric members. Siemg up the awboa ose merely modmn the failure mode.
Oi
-uxs
4.U.13 Moalub ad
MageMum b fal inAt disfavor compared to aumaaim for cam and houuap becamus shorWting of srvtice life due to corrosion has hewn. a sowfia ant miteneam and spam replcmet aepes to the Army. Be•muse asaaeiim is a. tc with susect to all other mmals. failuet to a•ploy Oade design. prmws, and pmventive mainUance mmsume has cued an ovrunplhsis of the deicin of the mata. bare magsum acsc4 is afecmted ka by exposure to msrine axmaphe &hfu il umaprONLmd mild steel (Rtf. 101). However. corrmui of smeeiua allos can be
4
::•; -.
•
ii:•,-..
-r --- ''•.1•°
avoided succesefully only if the designer and fabricater follow the complete sequence of I. Design 2. Cleaning 3. Chromating or anodic film application 4. Surface saling or impregnation S. Painting 6. Asembly 7. Routine preventive maintenance. The nmot frequently occurring inadequacies in recent Army experience involve design and maintenance, Aluminum alloys should be used in areas of high susceptibility to corrosion. Ref. 46 reports the replacement rate for AZ91 magnesium main trmsma ion ca at UH-i ovrhaul as: Top case 16.0% Main cane 1.7% Support cas -2.3% Sumpcase .% Quills 1.0% In the cane of replacements of the top case, 1/3 were attributed to improper protection of bare surfaces during stipment after removal of the main rotor mat, and 2/3 to in-seice corrosion. The relative re#piafemfen rates smeat that the environment in nal the top cae operates spartzrularly conducive to corrosion. Aluminum alloys with high silicon conttet (6 to 12%) have been found to be supo to other alumiawn alloys and mngnesium alloys with respect to wear reastance. Properly dosigned spines of these materials will exhibit negliwear when operating with floating steel mating spliuwes. Magawsea and aluminum alloys commonly us-d in hldicopt bousin and ca- are listed in Tabe 4-6. ,.i..-re.ina ad .,tneai kat he.. i ,c.nat,..,t4 aeT-", can asure the higest allowable stregth .preaties. Tw tuseof MIL-A-21l10 control -flcatio rmaher than t of QQ A-I generally ill insur 25% hither adiowable fatigue stenth, alwlumahl• ciee 251)6 w tUow fatiguersT~h,patenthough the cogtmay be20 to 5M greater. it tin weight uavings may be n high as 40% if s". etrength delons tk design msl provided that mini. tri.ion ae nt imposed. The ,mum wall Mt..it Semide propartue of many aluminum foging alloys may be iinproved by cold working or mechanical artr m (Ruf. 102). Procurement and proces specifications for asngand are defined in MIL.C40l21.T caslins da~nd forgiu frg aSm dn in MI r eL- 1.iThe detail doesgn drawing t require t h f(lTowin Sprocesnmd a mi(NnD)u- ts Ic
)wbkb
-ibg
i. lmproqnsion. Tlermo• ttuig polyester mins \pg MIL-STD-276 wre omomended fo casting im-
prqgnation. Vacuum processing isesntial to remove gas bubbles from casting pores and to permit good resin permeation. Leak checks may be acrostatic or hydrostatic, although the former is preferred for sensitivity and cleanliness. 2. Radiographic inspection. Radiographic inspection is required in accordance with MIL-STD452. The detail drawing must call out x-ray views and should include a stress diagram to assist in determination of techniques and interpretation. to be employed. T(he x-ray technique should be able to resolve 2% of the thickness being examined. Film interpretation is based upon discontinuity gradations as defined in ASTM E-155. 3. Surface crack inspection. This must be , comaplished by fluorescent penetrant techniques as defined in MIL-l-6866 and MIL-i-25135. Inspection must be performed after forging, final heat treating or ajing. cold working, stress relieving, grinding, wdding, and maci~ning, but before polishing. tumbling, shot peening, plating. resin impregnation. or painting. 4. Hardness inspection. All castings and forging should he
chateSl for hardens by the standard SIn]
kg Brineil (or equivalent) method to ascertain that full heat treating and/or solution aging has been a%comnpled. This inspection must be performed prior to shot peening, plating, or painting. 4
1
Quls
External and internal quills frequently arc used to bouse a gear/bearing subassembly to facilitate moduhunmaintenance techniques and reduc the compkxiof the primary gearbox housing. Problems typ-cally encountered in helicopter applications are ociated with excessive wear or high tempeaature pcre of the housig bore that accepts the quill. The primary ctaie of wear is the ase of material combinatkms that permit differential thermal expansion with lomotsem becoming excesive at operating wtp ursCobnin*ha itmativte toupetures. Combinations that at elevated this problem while retaining mperatures allkeuv em of ambly at room temperatures. Hower, the tlmDml stress effects at -6erF also ma be conidmed (par. 4-2.3.1.1) to assur, that the material yield strKenth is not exceeded. Decause steel liners frequenly are used iP li•ht alhoy quills, the solution mom us be eatended to consider the efcts of four conceinric rit p with respect to their individual tolerans material Mrengts, fit-up, and thermal expensmo coefficients. Sealnt compnounds always hiould he is at extenal quifl-housing joints to prevent water entrapmen.
"4-67
¶
TANLE 4=4 rICOPT
ThAN&MISSON 11 CAS•E MATERIAlS AND APPLICA11ON DATA
MATERIAL DESIGNATION
PFOPERTIES.APPLICATION. AND RESTRICTIONS MOST FREQUENTLY USED AL CASTING. EXCELLENT CASTABILIlY. BESTCORROSION RESISTANCE, HIGHEST DUCTILITY. LOSES STRENGTH ABOVE 250" F. GOOD
A356 CAST
WEAR PROPERTIES, BEST CASTING FATIGUE STRENGIH
A357 CAST
SAME AS 356 BUT + 11% Stu, + 17% Sty
249 CAST
BE-T Stu AND Sty ABOVE 3E;3 0 F. POOR WEAR. Stu 20% ABOVE 357. CORROSIONAND FATIGUE PROPERTIES POORER THAN 357
z
SXA201.0
L
CAST
BEST Stu AND Sty BELOW 3500 F. CASTABILITY, WEAR,
(AMS 4229, KO-1)
AND FATIGUE ALL POORER THAN A357
2014 FORGED
MOST FREQUENTLY USED. FATIGUE1 Stu& St AND FORGEABILITY ALL GOOD.CORROSION GOOD, WEAR
_
1
GOOD BALLISTIC PROPERTIES. RELATIVE TO A 357 Stu UP 10%, SN DOWN 7%. FATIGUE STRENGTH LOWER, WEAR AND CORROSION SAME AS 249
224 CAST (AMS 4226) 5083 FORGED
_
_
POCR
4UJ3 FORGED
BEST WEAR PROPERTIES, FATIGUE LESS THAN 2014
AZ91 CAST
MOST FREQUENTLY USED. GOOD CASTABILI TY.
AZ92 CAST
Sty HIGHER THAN AZ91. OTHERWISE SIMILAR CASTABILITY. AVERAGE Stu AND Sty HIGHER THAN AZ92. EXCELLENT STRENGTH AT HIGH TEMPERATURES. RADIOGRAPHIC INSPECTION
-
LOSES STRENGTH ABOVE 250'6 F
-JEXCELLENT
ZE41A CAST _
_
_
_
QE22A CAST
DIFFICULT
BEST HIGH TEMPERATURE PROPERTIES RELATIVE TO AZ91. St, UP 9 %npStv UP 6i,. EXCELLENT CASTABILITY. RADIOGWAPHIC INSPECTION DIFFICULT
51t - ULTIMATE TENSILE STRESS, psi Sty - YiE.D TENSILE STRESS, psi 4-2.4 SMCIAL CONSNDUkATIONS
Akboqh many special displines affect dai e-
Ss
m ru
t aniemce indicates that two of
pFm-MgMmqporia
an vibration conrol aW dia-
nioka. 4-4.1
i
C
di"s amum -enkumul
dymmis awe treatW in par.
5-5. AMNCP 70&-31. while addia conskderations of oin Soveraiq. owlap. smd dampig are dotaile in per. 8-7. This panqraah addresses probIsmi awmedap with componnt rucmaat vibration,
particularly as it affects urav membefs. Remnc
Army
helicopter RFP specifacatiom have gated that ea remnant fresunis
"m be a minimum of 30%
away from the design contiuou operaing speed.
Wbihi this may be impossible to adcieve with sonte Wars whem all vibration modes arc onsiere, the intout may be satilaid for the polatially davgerous modes. As an aernatimv suffacient dampin may be
oumpka
t tod -eader -uch
waomat froquencies harm-
loee i.e., the vibratory stressia will be sof" blow the amdiraam limaiw for the sructure. As aisag deuia saeds for drives have rcsuid is pich bee vs4otie above 10.0W) fpm. many
*[
J•
AMCP 706-202
frequencies readily may be identified. In many infatigu, failures occurred that initially were attribinstances, designs will exhibit very low vibration due to However, loading. tooth to dynan.ic uted the favorable mass and stiffness configurations of the aestigation revealed that the fatigue nucleations flange, web, and hub, and therefore will not produce usually were loes ad in the bottoms of the tooth roots a detectable strain gage output. Considerable input or (n the insides of the back-up rim. The crack energy may be required for a realistic determination. propagation gemnally was radial rather than acrtss Although actual operation in the transmission is the the tooth ban, resulting in the loss of a large sgtinal arbiter the acoustical siren generally will promeat of the ger rather than a single tooth. Such duct sufficient input to get the job done. failures are typical of resonant conditions in which Audible detection ah3o is sufficiently precise, althe tooth meshing frequency or one of its harmonics though a microphone feedback coupled with orthocoincides with a particular natural vibration mode of gonal axis input from the exciter into at, oscilloscope the gear. is required to produce Lissajou patterns in order to often use Lightweight gear designs for helicopter distinguish between fundamental and overtone will exhibit various types of vibratory modes, such as responses. The observed standing waves are the with radial nodes or circular nodes, singly and in product of a forward and backward traveling (with combinations. Typical vibration modes for a thin respect to rotational velocity) wavmset. If the gear is Ref. in given are web spur gear with integral shaft rotating at a given speed w,two resonant frequencies involving that is ofconcern mode 103. Generally, the are obnerved for each static fundamental radial vibrawaveaxial into rim gear the put that nodes radial tion mode: form vibration. The lower orders (say, up to the fifth Forward wave natural frequency diametral mode) are more likely to involve higher 4 - f. + w/2. Hz (4-M) amplitudes and, hence, higher oscillatory bending stresses. However, relative resonant response amplig* .ard wave natural lrenuencv tudes for constant forcing input intensity at various diresonant Lrq!.%eicr vary enormously with fsei Sear blank co•r•furations. One gear blank m-ty s2. W (4-5) -respond most to a third diametral mode frequency where while another o the fifth. The flange, web, and hub design, all influence this relationship as well as the static resonant frequency, Hz number of radial nodes ratio of higher order resonant frequenciea to the funfotitional speed, Hz darnensal. The resonant frequencies ma best deA graphic presentation of the phenomenon is contermined by experimental bench test techniques using tained in Fig. 4-34. The fundamental radial node only the gear in question. Excitatki can be by mitihanical shaker, acoustical siren, or electrostatic resonant frequcies are drsjgnatkd on the ordinate by the number of their radial nodes. The abnignealy by an induction coil mounted very n;ear WM i's r.tinnal aerea anre-Inliv tfn nnr..al prra" the rim surface. Excellent visual de-trmination of tin speed a The inclined lie represents the gear mesonant responm may be accomplisbhe with a"ed con-orI_: if the gear web is ofasuitable teichniques pattern freqecy for a 41-tooth pinion melhi bdlanino"da otheinsitanc~ebsfmodal fPsraaaon.~tt fiuratio:. ,I' otw in , nodal and antinodaf driving at a normal speed of 20,000 rpm. Note that maw r-.y tic clearly dettckJ by manual probi g of the forcing f'untior represented by the pinion tooth int the forward traveling 3 node vithe rim a•A- %-vbsurfae with a lightly hand held soft brion nesr ground idl spee, the backward wagv S1fl ly,. AAible detmctl 3n is also sufficiently precise meh -
3% speed, the forwardw0 fromts I0 node at about i fkw mmsamt f•quemcy identification, although a nord wave at ner normal opeating speed, and to1 microphone fedbacks coupled with a radially op. The 3, andtOW at ovespera poed imint fromn the exciter into an oscilkoscopewsrd 1isnode wave 12 node vibrations a- all penmtially hazardous. pod, between pftetalu nandary to required Even if it wer de.or.rated that the cyclic stresses disuimaiah between fumdayvental and overtone wer below the material enduran.cr limit, obMasngne segnawios mqay beased.e reeponac jeetsaable acoustical enrgy radiativa would occur feedback phaug sial with equal -e. Whaen wo vibato amphiudes arm sf- a thee iatercb. teSig of the shape and mats of "ficiumt to prodm wsiificant stress levels, te most the ger is not very practicable in this example bocamu a 25% clhang in ear rim and web thickpractsi qumaiative etvaltutim can be made by ataass affus maxiamur chag of 3%in resonant ta, bamstrain gag to appropriate antinode regigns frqiency for a specific deior- (Rdf. 10). f tshe rio anid web. The hige• emeI lvl rsunance
-
N
-
SPIRAL BEVEL GEAR
o to
to
9
a0
11 a
aX
go
IN
11)
DRIVE SY$TLUS. 'U -L 1hU LOflMAT.OSftEoE___
Gear Teoth Mebing Sped
duction of sufficient damping to vudmn the vbrmimo sigificantly. The spiral damper ring or snap ring a shown in Fig. 4-35 has promc to be efeti1ve in many inisaes,a producing damsping ratios of 0.04 tog 10t With a resultat reductiont of Cyclic resonant arcs to 30-40% of the undauped mugnitods. A*16m.r
r
-A-
=
--
SW
I
1
-__IUI
dA-P.O
otte ty, an of damping (viscolstc ard torsional absorbers hat: baen evaluated with pautial mance in bulampte tflnitia s. Certain vnecdmiet damping tremcaat have boen shown to reduce some Var vibrefion mode amphwudss by 50% (Re. 105). -hruhadcapeesv
"AJ Dipm~estotchnique Many cockpit indicators warning lqlns. and gpips lit ths postal deflimic.o of daagaswi aids. Recet Army REP specifiations icmii indications for oil presein and Impsrathne low pesusaa warning, high Itw~mqaeri. warming. aid quantimty, and chip dea1e90m. Atddmitoa grouned iampulae wfchanquu rutial "hinc Rd mpending Oil-rakertypass warnin Rw oil beaings chas viswa! and am;l faduke deaL~a and lacti, and oilsm.. hag for spectrgraphic analysIts. However. the n~ thats by theisles haeve prowcn inndsuguaee to allow the wiae*rod OV saf and cast-daffve conditiontal
4-70
________
p~ 43
rypsal Sis oeinp. King A"Undimsim.
mnumaafah on existing Army helicopter drive subsyitwus Tou aitpaextent. the mussing ingredient ms consideration of these in the initial design and development test phase. Safe and cost-effective inapismantation of condational maintenance methods requires thorough definition mwand n ate ume of early failurt detection. digagnosisad pro~wmoa as defined in Ref. 106. Majo effort underway in Army-sponsoed programs sam aimed as reducang the time required for deei*om and as immproving sceduling. This philosophy rewts in sebmlsedaisg of unscedul~ed maintenance provided that eqAp~mnus-bae judgmients are availtli aommarig the secrity and mi of progasic of t4e deated fatilur.
-
-
--
~
)
J
Detection methods may be classirted as related to the internal oil system or external to thc oil system. and include: I . Oil system dependent: a. Spectrographic oil analysis b. Electronic or ekLetro-optical oil monitoring c. Oil filter differential pressure d. In-line electrical resistance filter grids e. Electric chip detectors C.Magnetic chip collector plugs. 2. Independent of oil system. a. Naise analysis b. Vibration moititoring c. Temperature mma.urement. SpectRogrphic oil anslysis has pro~en expensive and relatively unreliable because poor correlation exists between faiiure severity and detection. Little is vacually known concerning failure, therefore. diagnosis and prognosis ame impossible. Oil monitoring rouains hr, the cvailaatiora stage-. some techniques havig proven totally useless while others show sowe promise. Oil filter differential pressure can be correlated quite wen wid, iaiiure progression rat;ei of bAriqp when the rate of chanae of differential pressur is examined,. Usefulness of this system iscomprcmis~d. however, by the accumulation of normal wear, dirt, mad contamination parlicbs. Visual inspection of !he filter elemnwt debris by t. perienced tochniciantn can detet mine ahe generic failu": mode buta not the location. F~ilter grids arc best cmnpoyeJ cc a Wahion similar to use of oil monitodriat. The hasic diffacrice is ttal tnefriter grid paitjaz sw~Albr con..an~inant particles. hopgfully rstainiivi, Only the laiýgcr metal fudturc flakes. The of.s~ @1ri cioggaig is detected and may be coVCtvete Cteri(Mlly to Itrae-of-chiingC display. Chip detectionis have bftc used extensively in Army helirvpiors for two dftades. The debris parlidles ame atpturod a~agaetktly to bri~c electrical cnotmc points, which, in turn, ensrtize a caution fipt. The wsduincns of this metho is dependent upon the locaition of tkm dated". 'siA W ~ resulting crew action. failure indication often rest~ts from thr sdow accumulatioa of miormaa wear debris. and oftencn in awiteraw t oftn roosi te cew reult ina itisio, aortto 110vssttpaW the wCas. Use of chip colhctor phigsu eUhaiats~ an aIItl4A at eatr i t~ ie~~es ~iway lawi.nsmbmof plstwith greater deris starlac capacity and W. teual caution panel light aft used to aid in locaiMinof a fiailjr to a specific area or module. liton is aocoapqlished at daly oi Meiodi&c inte-
~A"C 705-202
vals and sufficient debris callcction isusually prc,.nt to permit an expericnced technician to identify the failure mode. Good sucmes has been claimed for such a system on commercial air carrier fleets, but uc~cesa in helicpter adaptation is comipletely dependent uponl wcll-designed installations that entrap and localize failure debris, convenient and accessible delector loca'Jons, and experienced technicisrns to correlate findings with other available diagnostics and to schedule maintenance accurately. Vibratioat monitoring and noisc analysi~differ only ini the sensing techniques. Acceleromnetu or other forms of contacting vibration sensor% measure vibration, while microphone sensors measure noise. Substantial research and developmnent effort.s are being mad.; on these systems. Ruiiimentary go-no-go Syak ins with a wamring threshold signal reqtiiring aigine shutdown have long been used in turbine engine installations. The real challeng lies in the electronic signal processing and its conversion into an identification 'and quantitative assessment of the failure mode. Signal analysis is being investigated by such techniques as auto-correlations, Fourier transkfutal Ma51711, jUWFA fiPFtiss
dmbleftj caiaaatL.~.
aross powcr spectral density. amplitude probability distribution. anJ real time dlassificatirm of wav&form by convolution of other techniques (Ref. 107). Tht operational succes of any such rystam is dependent upon its ability to isolate the faulty comprpncnt signature from the background created by otW~~ internal and external forced and resonant vibratians, siske bar~Ak and best frequencies, at an earlicr time tihan the chip detectof, and to retnain onlint to monitor the rate of progression of failure of fth particular component. Onily in this manntr can the remaining useul safe life 'DC pradmad. and maintenianc scheduld wisely. The ultimatae goal of these programs is automatic identificatiott and prognosis as well an detection. This goal requires coanpmehensive test data on all failure modes, both individually and in combination, along with impravad sensor reliability and an analysis system havisv;, minimum reliability an order of timag *d grnLter than that of the drive system being monitored. Tepraacwsrnctsvm, intAa Temtue prctorelativel se wa madsytwrlmpteatur ul toercrtiel integrated bulk messurmncit, seldom cam detioa other than viny advex" (virtualy emominico conditic") faikies. Capbility imnprovxnent roqapinz &multitude of individual conipowats *one=r caxn~iwsd with triw.ofdu nlssnawdfrbs ln ai6ai cukhavife umiays creted (ora selnevriwoni Rapid improvemients iu &
wtset"~ for bvlaaogw 4-71
I
4
0A
3:
At
dfiiv' subsylensntam NWAutkf larEZ4s UPon: 1. Improved kiixia dcsi,;n ra.uilure-forgivitkg d4sigjn lchriqbt" as in ptr. -1.21.2I. b. Integration of nonssay Gdicstk ais& mW gsospr provisions into the oa~nx!~v *tiaWr. 2. Redireca edevelopment ttstirw effort a. iog out of failuremi~ as~d MdWscImd in par.4-14.43. b. Compalat of pwogrssicon rate data and correatio
lvel.gearbox, wi
3. Implemen tasiono orfie ao~an
f
efectuive
un operating
of saI
a.Deemntio t0t
umn for
cvrow
Wa y
ar b. Dlriaion of. ipctosc failure modt4s from exciessivc continued operation.
43DIEvs, INTWRCONNECiT SYMFMS In a halicopte trnuasmncui and dc~lr utri
PRh,
~
kW i,
is erWnx1csa.'d
t% X~l
fth.
.&flaygtrI, CflCI'i
thfa wu.4&t ac- a ge.ýrb*Ný mad f10o¶ a itiria
to a
k
shafting must a-ecohmotlatt. These mounting systems may be giou;-W conveniently into three
0v
If ý-.
categories
Engine(s) mounted dirccly to main gearbox 2. Enin mone to irfa with tubulr strut&,, usin rod-nd bearings to relieve thermal and load deflsctious, wish the pzwrbox bolted directly to the airfrmie
Engine moumbx as in Item 2; gearbox mounted flexibly. The rotor pylon usually is integral with the that is. airframe mounted with elastomcric3 springs and hiSWe liks or struts. Syrteni I generally usa an internal splined quill be ~~sha;'L lubricated with leaf box oil. Thezspliuics mnný. hardened and ground, slightly cio~ned of straiZln: or indium hard and bobtstd, shaped, or roll-foamedt. yso Wa n rtn orenoyL tu* pouting and/oir sof, plating or coatings such as xiinickel, nylon, Teflon, or racolybdesun d"isf~c vhce Inva~a~sarcqpiýazsacuta for Systemn 2 can bc moc&
-".? k
W.psrdi
s3 m u~nd wu[!ciiAirgi- with rralaivcly$k arith citteve& Aaft1 Ul., retc.¶vrI tac
t~
rur'dMg*e
rtc"-n I 40r'r
ThOC 606010a *AraesI. W to pCci&ubCdyntamic axial defecion b vatiatiozas isctakic kugibh due to installauo-r. coj~~~.bearinsr, and benisl wukmt (hangusý 44,3CE~i~tA tL~J~a~Cmi&i;iwbeflL06)O tx4 x;tho&Si'T growth of the engine must tc &lcdticuor V-1mSs~uatW
Thia ncuvia4 darv w2ita tgks, Id 01prupsLa. t
¾
SjLa&Ii driveL uhaft rsqtirwaco M an 41 &.umin4 rv-*iwa~j kurn~od.by Sp.Aizdrveof m ,b.tparticular heicopter- dc4A.ý. iincu&ni thet &WA.t wzseý.' cuup.a. hagemiw vtgiu slce. w dftz,. and otAýt' VoJzlh)iiftio 6cweit #i. GO~I "P,
t¶haft dzsign. Car must be taken to stiructure provides support at wmtua te t Ota ew~~ii and gcarbx attach points adequate to ddfecion. It often is possible to comply ith ujhkzut-#ad cra'al loci I2IwtDZ Ittjuircnents.
pvamasily by the poi-va scqjamnsmee.L of Otnt,~i llo~ion to b: cactaxI vu ttt.Zity. mwvtIa~~imnbiPjA~y. %4 ik. 4 amuaAuoL "Iat r~&1% w.4da. ivw*fmnaeuta for relinbihiy. maiuwe discumrJ in par. 4t tnio.64Ni y~.m swuvability
)%AEVtM4Wr rctL'iVdy hilh dcfleciMý mjoA4nOrU# aitfrauW nw&I. hi. jgivcst loi tr4~-w;. v=OV sad Ag4udc&~tirww of I dtjý sa k~is nky arc org-
4Gc:%Sai
1.bM
Wdz t Lbagt all-iacuuimc ncWjuiraerw am, tswtrA pe641C"kb vahal.duab"'W iue
In
to
ftlrewijuurtmt
O U4NAIMC w ang
q,"dksc thi~t kuaksm aw-4dwcast in I~c pwar-
cow&aLUt4 byZ~il did: (Thotuam type), r"i-A 49g l mxicn tentoson ck dizra--st (ILandira t 4ixaq,menz Jbalaher tyipto,wcrtnwm4 tý.'kA gear ccupiw'L."ý The first 6.-&typ" SPw.a&Ny asf. preferredl tnx45iwc
t* &ru
Sysan YconVW~t
J
X1C.W;
d ni k~~ix*4. liaidt i~ aflU 4~fa.n~~aatat~tpaa~ysaa~n
Idnh ia.
M0ealt"o
r.ti
Vd~A,0 W..'4r .fltt*
catar
oa
(Sol
ottstk%1=44 dndbcntkWC in 014f
Stt"ic
balk~ou
ianjWkvt tl*& 0-&
Iubrication.
)144&4 as
70
&xW twý2v'az uŽ±i~u ~,sk
'
~
1
4ý
The an junt of axial motion that must be acas functional failure of the shafting becoomes a catascommiodated usually will determine the type of tdrivc trophic Walure almost immediately, with acollison of shaft couplings used. Alt known drive shaft sysniý;s the intermeshing rotoi: blades. Achievement of the offer a resistance or damping force opposite to the dinecessary level of reliability requires detailed conrcetion of axial motion while triasmitting torque. sideration of operating xtresses and margins of safeThe maximum acceptable value for such forces usualty, critical speed margins, number and type of dyly is established by either the rotor vibratcion isonamic components (such as bearings hangers. lation system or the engine PTO desin specification dampers. couplings. and splines). rodundari'.y in limits. Damping force characteristics for various mountaiag and support itni~ctuMe and easw of incouplings and spline combinatio~ns are discussed in spection. Criticality of the interconnect system allows-f par. 4-3.2. 1. little latitude for rchal'ility trade-offs and2 comproOther input driveshaft design criteria are governed wise with weight, cost, and m~Aintainability goash. by maintainability, v~ulnerability. and reliability Optimization of design. then, must be in the direquirements and by additional engine pTO design rection of minimum number CL parts, low stress (high and specification limits, margin of safety), and Whb tolerance to ballistx Maintainability considerations require that the damage. Therefore the drive shaft tubas will be rc-laengine-so-gearbox shaft contain "quick-disconncct" tively large diameter, thin wall, and long (within a featur s. Ease of accessibility also is required to facilisafe buckling) length/diameter (LID) ratio and crititate drive shaft inspection and servicing, and enigiric cal speed limit. Intermediate bearing hangter design or gearbox repi xcement. Since these tasks must be must permit relubrication, with ready access to the performed at the direct support level. the absolute whok, hanger fot- visual inspection. The selection of minimum of special tools, fixtures, and skills should drive shaft tube material can noxuitate further conbe required. siderations of a"Ia motion due to differential expansion between the aifnframe (generally aluminum) Although vulnerability and reliability have been dtbmUsseJ nreviousiv.i is irritortant to co~nzio n thec 1104wtfvesbait (Step], aminu tivaniuamo comsequenee- of drive shaft failure. The large kinetic composite). Airframe deflections due to flipht energy of the input driv shaft categorizes it as a maneuvers; or load distribution also can contribute to potenitialfly hazardous o. lethal okiect should it the axial de~ections of the drive shaft. These deseparate at eithe.-or both oF the engines and gearbox Ahctins will necessitate couplings capable of abadapter. The ueof auitiflail devices, i.e., secondary sorbing the anticipated motion. if axial dameciotsi components -or structurt capable of capturing a failed ame small, then flexible disk couplings frequently are drive shaft. is highly desirable, the choice; for larger axial deflections, the geared In addition to limits on &xial force specification at coupling or ball-apline disk combinations are Watter the engine PTO pad. allowable moment and steady uWted. Under any specific set of requirements, the and osiLlatory radial Icads are usually slocrrned. primary 4esig empLasis must be reliability and Sinmc egine-to-geautox shaft rotation speeds are in ready-access fra sevice and inspection. Ow. rang 6W-20.GOO rpmr, compliance with the osciilawory lod limits generally require kincoatic and 4-W u u Satn dynaniic balancing of the individual elemets of the ~ satsse ~do igemi oo ThdrvsafsytmLadoasigeanrtr "trve shaft assmbly. When positioningt or locating helicopter to power the Ladl or antitorque rowo. "mai tolerano. between mating surfaces or elemcnts suc, couplings. adapters, and ahIts cannot be conbeft-een the main gearbox and the tWi rotor gparboa. This system smSt provide power to the taul rokir ay ecot.- ~ccsar to tro~d aequt~y.it masn y toelominatessarym th mi rotor in the event of loss of drive from cgitd trolle adeqaey balnceth mialy covleeo inuateexessve thme engine4s) In ormal operatio, the en~gine*s) vibration.driv thrnoug &fraewheeling dhs-* to tie main yewibox. During autorokatiom. when the fre-whselir4 "411 EUdim and Shellin unit is overrususing tW1 rotor power is extracted from An interconnec shaft syw .n for mul, Ole main (or the siam rGto adlorotAticnal, or kinetic flyweeL.iq.% liftin) rotot Wkdhpters transmits power between the artia. engine Vearbox (or the collector gearbox in multiTail rowo drive shafing win be subjacted to severe eq~sa eLi6copters) and the main rotor goarbox(es) trnasient loads and cyclic tersional ouciasioos as while also maintainin phase reiationshp between wagl as noual Moeady torqe inputs. Torque require rolors. The pitaary comuiduatioas for sisca an intermuass for moat flight conditions awe maderia is 000toc sha4 s"ste ane reliability amd survivability, nawne wit maxismum steody torqu requird during
4 I
-..
'A%
Ias
'
4-73
'r~
Sa - V.~
4
bover at high grow weight. The total power roquired to he ver ismain rotor power plus tail rotor power requkied to offset the main rotor torque, plus losses, r.A taW rotor also must counteract the main rotor cy,l4ic thrust vactor and an aerodynamic drag couple from the taillboomn. Conventic-sal rotor or propeller theory. including an effacieca factor applicable to A*. specific tail rotor can be used to, calculate the Mmady tai rotor torque. However, experience has shown that the transient torque requirements can be from 200-400% of the steady-state design torque. High levels of transient torque result from sideward flight in an advems quartering wind, from yaw acaelerations, and from unusual inik , conditions ivsulting from combinations of main rotor downwash, tail rotor blanking from aircraft structure, and adverse winds at hover or low flight speed. Transient torque inputs also cen be introduced to the tail rotor drive system by engine compressor stall, violent flight maneuvers rapid throttle movements (chops). or abrupt engine power loss. Under such conditions the abrupt relief of the windup of the tail r .4or drive-shaft combines with the flywheel inertia of the tail rotor and with secondarv effects of main ru~ui iwlin-at to cause aeveral cycles of extremely high amplitude torque oscillations in the drive shafting. Although occurring infrequently, this low cycle-high stress phenomenon can cause fatigue damage to thc tai rotor drive system unless the components of this system are dlesigned for torsional loads well in excess of the noirmal steady powter re~uircments. Transient design ctitcria for the tail rotor drive itipulated in MIL-T-595S and AMCP 706-203 are 300% of the power required to hover at design gross weight and denisity altitude or 150% of the maximum power rcq11irnd
in th~meet sver,. mzne-otuer .;.th;n
fi
infinite life critria. The throttle chop transient response is often the greatest oscillatory torque felt in the T.R. drive syrteni. The system can be modeled for the computer iiiring the cnginc-main rotor decay curves, the appropriate lumped mass and spring rate analogues, and the coupling discontinuities. with reasonable accurascy. A prcproduction flight strain survey will provide sufficient infornmation on the torsional charsctcris-ýics of the tail rotor drivesystcm to enable substantiation of the integrity or revelation of the unanticipated weak points. 4-3.1.4 Suherltlesil Shaffing Analytical methods for determining critical specds of a drive shaft are covered in Chapter 7, AMCP 706201. As defined there, the critical speed is that rotational speed at which the elastic forces arc overcome by the unbalanced centrifugal forces and the "bow" of the shaft increases divergently. Theoretically, the critical speed of aperfect shaft, i.e., a shaft that is perfectly balanced, homogeneous, atid equally displaced about the rotating axis, will occur as predicted by* analysis. The behavior of such a shaft is . dashied line in Fig 4-16. WN vihrstirnn shown occurs the rotational speed 17approaches the critical speed n,, where divergence occurs almost without warning. Practically any rc I shaft has some initial unbalance that provides a centrifugal driving force which increases with increasing rotational speed np. Suich a shaft exhibits vibration /rotation characteristics such as arc shown by the solid line in Fig. 436. While vibration levels at normal operating speeds
flight
envelope, whichever is higher. Such requirements are rather straightforward withFISCITC[F'5)
respect to fatigue design of the gear teeth and the can., tilever rotor shaft in that the need for infinite life criteris due to the high rate of cyclic accumulation (rotation speed) is evident. However, with respect to the remainder of the drive system, where start-stop{ cycles, throttle chops (T.R. inertia overruns), airburm engine restarts. and yaw control pedal excurvious accowad for the bulk of the high stress cycles,.____ a far lower frequenicy of accumiulation exists. In such instimice, past experience iv ith the fitting of theoretical spectrum analysis to subsequent flight strain survcy results is rather essential in efficient design work. When fatigue spectra are unknown, a ,
rewaable appiroach has been to design static yield
strengt lev&l to a minimnum of 3 times thc transient fatigue stress used for the gear teeth and rotor shaft 4-74
~
R*F [C 5 fVSHAI TyrCAL SHAF1
-I
II
,
~..~.I_____ tATSAF'PE
*,,,
Figue 4-36. Relative Shaft Speed vs Relatlhe Vibration AmplItude
V P" may be acceptable. the unbalanced forces can increase rapidly as i increases above 9, with the possibility of resultant damage. This situation can effectively reduce the critical spead margin to an unacceptably low level. The inference is that simply by balancing the diive shaft an acceptable critical speed margin easily can be realized. However, the ,ast of dynamically balancing the bhat and/or shaft assembly must be included in the trade-off, together with a careful assessment of the contributions of the end conditions and/or mounting compliance to the vibration/rotation characteristic, The majority of existing drive system applications use subcritical shafting, for which the lowest value of It,, > V,. Requirements for balancing can be met with ordinary balancing techniques and equipment; relatively short shafts minimize production and logistic problems; and ballistic tolerance design paramcters arc known. On the other hand the cost of a subcritical shaft installation with seeral separate spans
BEARING (1 SEAL)
may be higher than that of a comparable supercritical installation. The i:,anufacturing cost fot the short shafts may not be much different than the cost of a single long shaft, while the number, an-' hence cost, of machined parts probably will be higher for the subcritical installation. A single span of the subcritical system consists of a drive shaft tube with end fittings, drive adap'cr, hanger assembly with bearing. splined adapter, and coupling. A typical example is shown in Fig. 4-37. Design of the drive shaft requires a determination of the shaft cross section necessary to accept safely the steady and transient loads stipulated in tic pertinent design specifivation, and of a shaft length that will operate safely within the critical speed limitations. An efficient design generally consists of the least number of spans with acceptable critical speed margins and torsional buckling strength. Large diameter thin-walled tubes, generally of nonferrous metals; a greaw lubricated bearing sealed on one side
BOLTED JOINT ADAPTER
FLEXIBLE COUPLING
TUBULAR SHAFT
GES
HANGER SHAFT
NUT
ADAPTER 3llf
706-202
//
_aHANGER
ASSEMBLY
S......MOUNTING BOLTS TUBULAR SHAFT
\
TAIL BOOM STRUCTURE
Filme 4-37. Typical batrilgl Haiser AMedy -- Sedicfical Shaft Ase-mmy \
4-75
with a fitting for a periodic relubrication: and a flexi-
shafting will be partially or totally offset by the addi-
ble disk coupling arc typical of current design practice. The end fittings and their attachment represent a considerable portion of the cost of the manufacturing of such d'ii-a shafts. The fittings may be attached by adhesive bonding, riveting, boiling, clectron beam welding, or brazing. Tolerance of mating pairts must be closely maintained to ensure good parallelism of end fittings and low vibration characteristics. The type of ccuplings selected, their mass. Io"cation, and friction characteristics, influcnce critical whirling modes as well as torsional response modes ofthe shafting. A recent investigation of coupling nduccd whirl phenomena on turboshaft powered helicopters is given in Ref. 10S.
tion of a daniper or dampers. The elimination of hangers and the attendant maintenance requirements also may be offset by the addition of maintenance requirements for the dampers. Acceptable tolerance to ballistic strikes requirs hardware testing under simulated service conditions. Parameters for ballistic-tolerant designs for supercritical shafting have not been defined and dependence on individual tests is almost complete. A method for calculation of critical speeds and bending modes for high-speed shafting is well presented and explained in Ref. 110.
4-3.1.5 Supercritical Shafting Supercritical shafting usually opc. "-s at a mced between the first and second critical specu of rotation, although even high orders are possible. "The main rotor and tail rotor shafts, or masis, often pass t hkm
lo
h
•
fi e t
PrituCjti' .
snrp
h cfýn r
d
p r
opr h in n
operating speed. However, the critical speed is ,elatively low, the dwell tin-. is momentary, and aerodynamic damping forces are quite I.. ge. On the other hand. interc.onnect drive shafting and tail rotor drive shafting generally operate at relatively high speed with very little inherent damping. The advantafes of a super.;'itical shaft design aic the smaller number of detail parts and bearing hanger assemblies. The disadvantages are the need for dampers, which for reliability should be redundant, and the...... physical ~length of the"'• !hafts, im"-I:*-" ~ other .. . which ,,t• "•-may "- r"-" IRO"•*
VBl
IRtOKOl1t%,3.
/'%13V.
9
.Q1 201
6
13#0
%-41J11 10.%,
tars i'ay dermine shaft diameter and wall thickhe~s. Shaft sizes larger than thosc required by the power requirements m4N be necessary to maintain a LID rptio sufficient to avoid critical torsional buckling, or to counter a ;pecific ballistic threat. Once si.e has been determined, the design requirements for shafts in the supercritical speed range center primarily on damping und dynamic balancing.As shown in Re.f. 109, the ncessity for balancing to a very clo tolerance over the entire span er the supercritical shaft is paramount for successful operatio,,
I
4-3.2 COMPONENT DESIGN The basic drive shaft system components, coupa.gs, bearings, and shafts are discussed separately in the paragraphs that follow.
4-3.2.1 Couplings The primary purpose of the shalt coupling is to p r
v d.
..
4 'f
n
g ut|
- m
i s aah gnm
en t
an
di
x'; al
motion between various shafting elements and the engines and gearboxes. This relieves stresses in the shafting, bearing, gearbox, and engine components induced by bending moments and axial forces. The rcla!ive motions between these components may be due to airframe structural deflections. thermal expansion. or pylon excursions required by rotor vibration isolation schemes. There are six major types of couplings that have been used in helicoptcrs, and the selection of one among them for a given application depends p .. .greatly . ....upon the required dis.olrt ni paccimicrts &no thle ioaos that can be tirid
i
their supporting elements. The six are: I. Laminated flexible disk couplings (Thomas type). This type of coupling, shown in Fig. 4-37, is probably the simplest design for angulai misalignments I deg. It has been used oi the CH-47 synchronizing 5haft and on the OH-58 tail rotor drive system. Eachi driving sp;der may have two or three at-
It is incorrect to assume that a supercritical shaft-
taching points (four or six equally spaced holes in the oisk complement). rhe larger number is preferred from the v:ewpcints of vulnerability and survivability. The laminated disks are generally circular rings, although square and hexagonal shapes have
ing system will automatically weigh less than a subcritical system. Dircwtly comparable designs for a given helicopter application have to be made and the total instalted weights dctc'mi.ed accurately and compared. The weight saving apparently achieved by elinminating the hangers necessary for the subcritical
bcen ustd. One problem that has been encountered is disk fretting at the bolt attachment. This coupling features high torque capacity, light%eight, simplicity, and constant angular velocity. The torque capacity can be varied easily by the addition or deletion of laminates. However. increasing the
4-76
number of Isminatas reduces the angular misalign.
WELD JOINTS
ment capability or the particular dcaign. An ad-
ditional rclationship exicfls between the numbenr of &t-
tachmieat puiats and thc- torque capacity, and misalignment capability. A four-pax'nt attachnent (two-bolt shaft adapte;) provid.es the maximutm misalagameilt capability and Odso is the least expersive to mnufaclare. Thc flcxi~'le disk is capable of small axial drilicctions, and where predictted axiWa motions
LXEDIPRG ___
P-e low, this cotupling servu well. No lubrication is sliding involue spline to i-Zoomicouatr
V&rlM1jO1Fl; ;,____
initial shaft aw~mb!y length due 'o accumulation of
mrnif~triig~lcrancec. b-lowcicr, &.'x.n amial de-
-E'ALIGNING MONOBALL
i%4fion occurs tindcr operitting torluc, the dlip rt;-
sistanct of Itbe spline~ is so great thsit apprecia~ble axial .'orcc will be rploliceJ to the disk laminates. For splines of 0.is type the brealhaway slip force Fju rarely is less k;, -
j
.4 Q/D,..lb
(4-66)
-
Q ~-tcrque, lb-in. =pitch
diamttei, in.
In some special cases whc~re certain dry Nali lubri(.atiteris are applied to the splines (Ref. 93) break-
away forces of half this value masy be realized. 2.leibrdiap' ,ai (Scndix tyre) couplings. Thibcoulin (..e fg. -3~))~s boa ~e onthe OH-6 hel~copter. Thi, type couplir~g is gencrally capale f arnaimim ~aula mialinmci ~ ahout idey per diaphragm pair, burt very little, axial def~tao. ad~apai~iy thtefre ~a~st e ~axial a sl~1i~ Recrcultin i' i breaawa splipe i~ o it ih wiTC ser.CIes v snip uisbrekawy -agdmt stac pity ar ge'~ll force. F6, A~sj, the di~aphiag tc 5v~ snitv stoe faigue w~lueakt tocme oslltheytax ial eloftensol caai ar.d provide r.waly all of the axial deflection. ay e ncesaryto roUndkrm;;hconitin.%it or saybeneesarytype pro-i onuai coifaligin vi ae atideia imcnf- ige ex r centcbll(roftesimilar typ)baring at he n~cri1e~f~urcceterof hestak ~ tanscr %xialloids into the s-ipporting adapter and tofoc the ball spline to move. Although theoretically it i5 possible t&ýobtatin v~~'y low breakaway slip forces wiih the ball s!apractical considerations with, itspev-t to aainirr.ium length of the ball track groo',ea, vh'ntbevt of balls. and stAl p,!cvisions usually limit. 1hese rorce toa a r;aininium of .,
..
J's.- 0. 15 Q/D. lb
(4-67)
DRIVE FLANGE
Figure 4-3b. iiictl~ Dispbragai Coauplham All too often, the vibratory forces due to imbalance and r.ý;or vibrations lead to false brinelling of the ball track grooves, which results in turn in much higher forces w'Ah increasing service time. Althourph the flexible dianhraffm elemouts need no lubrication, the splines and the monoball occasionzily require lubrication 3 Axial~y loaded straight element flexi'hle coupling (Bossier coupling). This coupling (Fig. 4-39) requirts no lubr~cation and has the abiiity to accommoda,.e combirned axial motion, m-isalignment, and torque. A series stack of warped rectangular plates with a vertraliy located reczangular cutout leaving slender sides chatacteeiz*5 the coupling. Opposed corners of these plates are b~olted to adjacent elements and to end fittings or adapters. Design characteristics o'hscoup'ing arc defined in Ref. Ill1. Apiain odt hv enkgl xai m-.,nta; with some flight time accumulated on the H; -2 h~cpc (Ref. 112) and UH-I- helicopter. Thec atngular misalgnament cipability appears to be about 0.5 deg per ,-late elemen:. However, an increrse in the nu~nber of ýesments usW~ resttll in a teductinku in first 4-77
A
--
whirling aitic• ]s apad. A satisfactory lightweight design for an engiae-to-gsasbox shaft for mod•a•ac angle (of the order 2.5 des) 0.25 in. scillatory axial motion probably would be requird to operate in t super-critical rune if tte erine ohtput speed were above 600 rpm. 4. Elastomeric couplings. Cbouideeabl developme- t work culuinatiag with eaperimatl dight tasting on helicopuers ai s the YH-51 have been accomplished with this Vypc of coUplling (FiS. 4-40). FHowevr, all sucessful applications lhav had low angular misalignment and axA deflection requiremsnts. Efforts to develop higher capabilities (up to 2.5 dqg steady misalignment and 0.25 in. oacillatory axial displacmcmet) have met with failure. The low angle configurations LAw us•d simple rubýber ekmenut in shear or compreasiou. while for h*i _angles very thin, multiple layer, ruber-ametal-rubber combinations, such as are now co•t•ion in certain rotor system bearings Itait ben ustd. ibe principal development problens bzve boee customer fatilue d, -o r•ersed louding (alternating tension/oonpr.,Ao ) at high sagle/low torque conditcr,. The basic advantages of cbetomeric couplings amr his'" o.upiihria- (low shock an.d oie -.............
!ARD
IARIOeECA&ULk PLkTL•
Rfor
-
DRIvEsr'Fr
_-yokes. DRIVEsHrI hEM 'CU•uAREAusWed
COUPLwG ,ADAPTE
Figure 4-39. Bosser Coupling
7
.-u
i
no lubricaio required. no susacfuihuty to fretting corrosion, and polential savings of cost and mnintenance. lanei disadvatages am dereAioralion in an oiy environmint and aging and reduced ctritcl spaed dvi to high compliance. 5. Hooke's joint. The Hooke's or Cardan type of universa joint coupling (Fig. 4"4) is capable of mlativcly •Aigh angular mimaigi•ret, of the order of 30 deg at wadete spoes and I1 do a, h*ig speeds. Howeve, unlike all other couplings discussed in this paragraph. the output is not a cons.ant angular velocity, and significant bending moments arc induced in the attaching adapters and supporting structum.ý. Consequently, this type of coupling generally is cmployed as phase-matched pairs to cancel the cvsillatory angular velocity or singularly with systems that am very soft torsionally and hence can absorb the angular velocity oscillation. The H- 13 tail rotor drive system is an elxample of the latter type of application. These couplings have no axial motion capability and normally arc used in series with either a sliding (involute or square taoth) or a rec'-culating ball spinc. The input and output yokes of the coupling
at..h.. to th cros with cupped nidl-c bearings. When that co..,oncnts arc sized properly oscillation. angular and force givcl torque thel absivkaway sliding• of velocity tlhe adj-Afrnt spline Usually is wel within the axial load capacity of the coupling. Common failurr modze, re spalling of the "cup end needle bering. and fatigue fracture of the The needlt býarings require lubrication. ti. Gear Wtouplingl. Gtar couplings (Fig. 4-42) with higu~ly crowned enternal involute gear teeth :...ting with st'aight twoiaed intcrnal gca2 te,:? have been on helicoot, r.; far more extensive!v than all oiliAoter csmpling typts combined. 1 '.Xsc coo j~iugs arc capable of providirzi moder3tely high anguLr misaligninernt and axial motion during operation at high speed and tot'qur
Ž-tr
v,-ry low weight. Common
opfratin• • 'onditirci &:e i deg continuous and 6 deg ADAPTERRETAINER CUP
k
. ...B[A
SRI,4
N,,
"" I
A
SVAr T
ý,,Ai 1 R1 TAAINRCQu
SDrfIVE
Figure 4-40. listomwek Copling 4-78 I *
DIV
CROSS
ligle 4-41. $o.4e's J-•la (Umimal'il
.
-7
4
"ADAPTER hOUNTING BOLT CIRCLL •'•
INTERNAL STRAIGHT TOOTH COUPLING
BOOT SEAL
GREASE CROWNED COUPLING
Figure 4-42. Gear Couplng
transient 2nd 10.75 in. oscillator; axial motion at frequencies in the range of 10 Hz. Unlike straight o.
spliies o~r recirculating ball splines. the breakaway blidinst
force F,. cait be very low; actwtlly F, reduces
-..- _-
~
as the mi'aa~ignment angle increases. A comparison of th;, force ~o, a typical gear couplin., n.shown in Fig._____I_ ccn ~act becween loaded teeth is at r.*high sliding velocity r~. t; the angular miselignments. Conse, quenly, a stmerimposed axial motion is resisted only by tht rela, w'vely low dynamic friction coefficient rathaer than d static value. G.,-r coupling operating limits are thermal rate than 6i1.3up,- w~hich is the limiting contideration for the ri .- ,oupling typej previously discussed. Specially developMr gresse lubricants have provided the best load carrying (least friction) capability for gear2 couplings. However, the operating environment (high
etrifulal field, high mechanical stroking fire-
qucncy, *ý--d tlevatiA tempcrature) combine to make the -erst ma~jorit) of grrase unsuitable for this app'iitMiss.
-
___
_
n -. a iuiy M I SAIL SPL!III BALL SPLINE A-. *-%-fCftULMEDN
I
___
W -1
-j
-
IR_____
____
IAIUNTAGL e
Figre4.3 Bma'kapay SliWG Foresei NalgmC foe '/arlkm Spiae Dekvkn 4-79
A.side from the need for periodic relubricagion (60D-hr intervals ate common) the greatest difficulty with gear couplings is providing adequate sealing for the grease. Guillotine slider seals and elastomeric boots are most often found in high angle applications while modified lip-type shafts seals can be used for low angle (
After loss of the lubricating oil in the grease by evaporation or migration, the common failure mode exhibited by grease lubricated hanger bearings is overheating, failure and expulsion of the cage. and finally, expulsion of balls. Severe shaft vibration, due to loss of centering provided by the bearing, or shaft failure may follow loss of balls. Degradation of the lubricant also is caused by entry of water or debris into the bearing. The means of sealing bearings provided by bearing manufacturers are generally inadequate to preclude a significant failure rate in the Army environment unless additional protection is provided. One such means is to enclose the drive shaft with a cover to exclude the bearing areas from the contaminating environment. Another simpler, but less effective, means is to install rotating slingers on each side of the bearings with closely controlled clearances at the slinger OD. This providL shie.d against the entry of water, debris, or cleaning fluids during helicopter washdown. Although an effective seal may be designed that will reliably assure reasonable bearing life (10002000 hr), a hanger that is designed to permit relubrication can greatly reduce hanger bearing replacement. Frequent introduction of a fresh charge of lubricant can revitalize and/or purge the old charge of contaminated and thickened grease. However. re-
4-3.2.2 Bearings The criteria for design of hanger bearings for drive shafting differ considerably from the normal power loaded bearings used in gearboxes. The loads P to which the hanger bearings are subjected are very light (C/P < < 10 where C is the capacity of the bearing for a lire of 10' cycles with 90% probability survival) and sizes are determined by the torque requirement of the shaft through the bearing. With high tensile strength heavy wall shafts used to reduce shaft outside diameter, a relatively small bore (light) series bearing can be used in the hanger. Bcaring mounting on the shaft should be closely controlled to assure true running and that internal clearances are adequate to prevent radial preloading under operating temperature differcntials. Grease lubrication normally is used, and sealed nonrelubricatable as well as relubricatable bearings may be used. The lack of adequate internal clearance is a common design error found in many existing hanger bearing designs. Considerable effort has been expended, as described in Ref. 114, to evaluate greases for hanger bearings. The grease most commonly used is M IL-G81322.
lubrication adds to the maintenance burden and the risk of servicing with an incorrect and unsuitable lubricant is everpresent, but most lubricants will provide satisfactory operation of the bearing for at least a short period. The selection of nonserviceable replaceable bearings or relubricatable designs is a trade-off involving many factors such as bearing cost, maintenance man-hours, re!iability, and survivability. The design of the hanger assembly must be such as to prevent inadvertent bearing overloads. Nominal bearing loads are limited to shaft weight and rotating unbalanced loads, neither of which should be detrimental. However, misinstallation of the hanger can introduce static angular misalignment between inner and outer rings (shaft to housing) causing a moment load to be imposed on the bearing. Although system compliance (hanger, shaft, and airframe) may preclude loads of sufficient magnitude to cause spalling fatigue, the bearing balls will skid as a result of contact angle reversal due to these moment loads. Such operation will cause cage distress and overheating with abbreviated service life. Adequate relief from angular misalignment, in the form of proper internal clearances and/or self-aligning outer ring mounting, must be provided in the hanger design.
4-80
AMCP 706-202 4-32.3
Sbmfthfg
Design of the drive shaft itself is concerned primarily with material, size, and end fitting selections. For high torque applications, where tube wall thickness permits, a spline or similar drive mechanism may be used to adapt the shaft to couplings or other drive components. With thin wall tubes, an adapter with a thicke ione.mboron thicker section must be attached to the tube to per couplingt
in loxv-spi:d applications where balance requirements are not stringet. Tube stock and bar stock, bored and completely machned, are used for higher speed application where straightness and true running are necessary to, meet close tolerance balancing requirements. Composite materials usually are fabricated by laminating epoxy preimpregnated carbon or filament at zero, 45 deg and 90 deg lay to the shaft axis and curing in an autoclave. The composite shaft has a very high strength to weight ratio but the
the section modulus is constant over the length of the bonded joint, the distribution of shear stress in the bond material will be even. If the joint not so designed and an abrupt change in sectionwere were encountered at the end of the fitting, modulus a differential
cost is considerably higher than for other materials. Balancing requirements are less stringent for the composite shaft due to the lower specific weight material, but machinable material should be added at approximately one third span positions to facilitate dynamic balancing when required. Large diameter (3.0 s OD) thin-watlled ahlminucetubes have demonstrat exce d len ballsti to leranc e toandi velocity, frod tumbled 7.62-mm bullets. Torque transmitting capability is somewhat reduced following a hit by this, type of projectile, but he vibration charac-
angle of twist would occur causing a severe shear
teristics are not affected adversely for subcritical
stress concentration in the bond material. The strength of such a joint would be considerably lower than intended. The fittings can be riveted effectively to the larger diameter drive shaft tubes with adequate margins of safety. The stress concentration effects normally associated with riveted joints must be taken into account in the design of this type of assembly. Bolted joint designs are similar to the riveted joints.' Welded joints can be made effectively when ferrous materials are used both for tube and adapter. Normal efficiency factors for welds must be used when sizing the joint for steady torsional load and the effects of a metallurgical "notch" or stress concentration must be included in the fatigue analysis. Brazed joints also are effective fo. some designs. Induction brazing is developed easily and is a cost-effective method. The heat affected zone in the brazed joint normally is tempered, and the torsional strength of the joint must be based on the minimum allowable strength of the tube or adapter in the tempered zones. Machining may be necessary subsequent to the attachment of the end fitting to provide parallel and concentric mounting surfaces so that the drive shafting runs true. Materials used for drive shafting include steel, aluminum, titanium, and nonmetallic composite structures. Steel shafting is used for engine-totransmission applications and other areas defined as fire-zones. Aluminum, titanium, and composite shafting are suitable for interconnect shafting and tail rotor drive shafting. Mill run tube stock can be used
shafting. Composite shafting exhibits ballistic tolerance to 7.62-mm bullets similar to that of aluminum shafting although the tolerance to lower velocity projectiles, impact of a dropped tool, or handling damage is considerably reduced.
Adapters may be .Idhesively bonded to thin wall tubes. The adapter joint must be proportioned properly to avoid excessive stress concentration at the bond interface. can be done by machining the end fitting bore This and shaft OD in a tapered or parabolic shape so that the angle of twist is constant. if
44
LUBRICATION SYSTEMS
A helicopter gearbox can be designed to meet load and speed requirements but the useful life of the gearbox is a direct function of the lubrication and cooling system. The amount of power loss as heat is governed by the design of the heat generating elements in the gearbox. The lubrication system assures attainment and maintenance of a minimum value of heat loss as well as minimum wear. The concurrent function of the lubrication system is to carry away heat. Heat transfer occurs between bearing outer rings and housings by conduction and from housings to atmosphere by convection. This mode of heat transfer is minimal compared to the heat transferred directly to the inside walls of the housings by the cascading oil, with convection again taking place. The t6ird means of heat rejection is by direct transfer to fo:ced air in an air/oil or, in rare cases, fuel/oil heat exchangers (oil coolers). During stabilized operation a balance is maintained between heat transfer by conduction/convection/radiation from the gearbox cases and the heat exchanger, if one is provided. Some gearboxes are designed for continuous operation without an external heat exchanger. In this case the surface area (external 4-81
706-20 wMC wetted area) provides adequate cooling margin, especially if forced air is directed across the gearbox. A somewhat different mode of heat transfer occurs in gearboxes that are grease lubricated. Gearboxes that are rease lubricated depend almost entirely upon the transfer of neat from the gears along the &haftto the bearings, through the bearings, and to the housings. A secondary flow of heat is provided by slowly migrating grease as agitation occurs but this is minimal compared to the direct conduction of heat to external gearbox walls through the shafts and bearings. Tests conducted on grease lubricated gearboxes using USAF MCG 68-83 grease (Refs. 37 and 115) indicate that grease migration is not significant, The lack of migration can be an advantage in meeting fail-rtafe operational requirements since little or no grease loss would be anticipated in the event of a ballistic strike in the housing. 4-4.1 OIL MANAGEMENT The delivery of oil from pump to filter to manifold and then to load points must be systemr-tic and deliberate to assure proper lubrication and cooling, Placement of the oil must be specific to prevent surging, foaming, and cavitation. As the used oil leaves the gear mesh and/or bearings, a natural gravitational flow path must be provided. Traps around rotating components can cause excessive churning and heat buildup, thus adding to the cooling burden. High speed gears can create vortices that will suspend large amounts of oil against thehousing around the gear. Excessive oil flow to gears and bearings can cause heat generation and buildup greater than the amount of heat coming from the loaded conjunctions. Therefore, controlled movement of the oil after egress from the rotating elements and heat generating points must be provided to allow the oil to find its way uninterrupted back to the sump. Close fitting shrouds around gears, and return lines from cavities between bearings and shaft seals provide effective means of preventing oil entrapment and excessive churning. Judicious placement of ribs and webs in the gearbox housings an4 ample provision for oil flow beneath or around the structure will help assure proper oil return. The pump inlet placement and arrangement must be considered carefully in the design of the pump, housing, and sump. Maintenance of a sufficient oil supply at altitude is directly affected by the volume and depth of oil at the oil pump inlet and the effect of flow constrictions into the inlet. If the return oil is hampered in getting to or through the oil inlet, cavitation and loss of oil pressure can ensue. 4-82
In splash lubricated gearboxes oil flow is more difficult to attain. However, because the primary function of the lubricant in this type of gearbox isto lubricate the gears and bearings sufficiently to riinimize
the heat generation, the amount of oil required at the friction points is minimal. Nevertheless, management of the oil is still critical to the adequacy of lubricating and cooling; provision must be made for oil to be delivered to each bearing, gear, and seal. Natural laws are employed tc acccmpiish this; centrifugal head, gravity feed, and dynamic pressure differentials can impart sufficient impetus to the oil to attain directed flows. Oil splashed to the inside of a rotating shaft can be caused to flow rontinuously through the shaft by tapering the bore from the oil "inlet" end to the outlet. The outlet can be at the end of the shaft whete return is accomplished by gravity flow through bearings or it can be through radial holes in the shaft, with centrifugal head forcing the oil into the bearings. Cooling (though minimal) also is provided by this Piow by ultimaze impingement of the warm oil onto gearbox interior walls. Agitation of this oil is primarily by gear members dipping into the oil sump and splashing the oil to the housing walls, bearings and gears, or to the inside of shafts. Auxiliary splashing can be accomplished by providing rotating dippers or slingers. Maximum cooling of the oil can be accomplished by the agitation and slinging action, but care m~ust be exercised to determine the maximum oil level that cao• be tolerated before churning losses override the cooling effect of the agitated oil. Grease lubricated gearboxes have a different set of operating characteristics. Although the high viscosity of grease provides good lubricating qualities, this high viscosity also prevents free migration inside the gearbox. As a result, thf. grease must be forced to remain in the bearing anM gear cavities, usua!ly by grease is thereby The inffges. means of shrouds and br 'aptured" around ee'h bearing, and the grease quantity must be such. as to assure an adequate supply around gears. The percent "fill" in the gearbox is critical, as it is with the oil lubricated gearbox, especially the minimum level inasmuch as successful lubrication of the gears is predicated on grease quantity as well as location. 4-4.1.1 Function The satisfactory fulfillment of the dual functions of cooling and lubricating requires that the design be approached systematically. Oil flow requirements should be determined and followed from pump outlet through the system and back to the pump outlet.
I
--
4
I
Ok
mWi 108 Psotida for IfrMn ea of 80 WU aNMAN NOW mpgdlsd hi suc* a way as to bured bum h~ ans4*o~l~ti, ed/r in.. pc.awd gu. The amount of M#t pmMM &P bei 4 oil dqmw to hk~d a ce~aueiom. i~e.. to pusV" to mat " aGu, isVM.,..a. 14mwge, if
r & pas. owm mnratream sow yti b &w r~ei. piam 9m IMI $MAm.md uofuSi Jai A sbmee of so*h sar q ios bA d ami Fig. 4. ads 1pm cam be provide by ao ao44.7m P me QWur &.'" PrWih p"M taheoff Ifro he male *1 ust. Wd i m dfmls miap tmhragWhd v the
INhiu m powibe al m ala. los
Ow interalci is &Th. S& dry~e."W6 o j nr abob.sa m" poy ofa o flodto m k~ shep6 fric-djbwdfs4tobaopadV toapetwb tpeboarand wMat &mty site pw ww = oils.Th sapytmf. wieh epetis.The eosda buiip eq'cman cma load ino ThslightymaeWr or gebothsesj awl are submctt WMa. 6"k hsak Wof c sh hadd.asst f Jthe dilin fuAcicatim eafl drah fight as epud. A lightly lade w ge.mit of euusiida, wit wihacpo nd elotrorcveipae WleMirop thbticasin wca m ert vcn W, rEse fne Anothe sac. for the dry spsytem is ith asysta sholam soat andw mll evuel vin htol cooingstabl ofgaba.Tedysm suiiigasnl yi urcation systan. eThe er n rmal eoegsro equtel by aacasne-avne /i oilde Sl".hames co4tc olubiae adkv mistor Ccnnsla y sl~h hghl lodedgoa mq~aneuvest or is p pwo sump ar ujc. t
:
d~id
i
.tS-jt,.a
mesh
n-r.
0
ahkto will generate osierbl hemidtrquief-a
~
ma? br.
-Id dqutl.
kirc
ýW
USG W.. RUS ova
flfR*.
yiji
A IW
man=c t
true For b;th;ik~ loadd and beavily loaded bearaws.flow 441.2.ndCwwsa; Aragemrmaw An1. oil 3 ")51cm aa idA num will co a s oI spoply of oil in a gei~rboA an~d it mewas of gagging oil quantity. e.g., sight glass, dip stick. The stamp must be &,a 4ocatecd that oil circulation vwill be accnipishd b a gar r rtatng lonc~t ippng nto the samnp and q.pashinu the oi! to~ the gear and &sarina ckaýients. This sari angt~men: for iplish lubrication can bz uLWc effectively, irgcr. botes with single mrshes not ins II operatir'g at light load where thc w~tted *1tr'bo arra providecs edcjuate couvrcctivc coo:IdLL. TMi sutor drive and acceswory d~ve gearboxtmL fel in this category. Although the power trainsmitted cyTb titjr cro cwioul slih th,, condition a, trwspstit and bulk hu*t buildup is jgtcui Lly negtig~bl. Gekrhox heat loss at hover powet can b.- Itasdircd cffvt1ively from the gear. box hoW-ni5wst the airflowi caused by the retor dowvnwash. Less power is requhcci during cruise car.ditiiwa, and rnoitt airflow is avail&ble. Large wetted areas at, 4; eunmjon for acccwry drivye gearboxes, and powcer requi!ei';'11s are ptedictabic a=4 exonsvhit. Thc sslari sysienr, i3 4n~ inhereotly we: sump 4ystenai. On 0vt othoer end of the spcwwni arc oil systtemsl K Consistieis of vil pump. oi0 lines and PWAwSC;. M40r av.filter. rlna.;Cold, rgUetof. oil os&Lei termal
ii
not i
mutbesavne
BUSM
nfrttdysm gfro ll P~6iYpmp in gni
ye
shasse
eoeg
win be OP tee
n
cofl-
stant disptacement type, simed for the. pressure and rafte determined by cooling requirements anid system pressure loon. Pumps can be desigd to widey varying flow requirements with single tanant pumps possible with flows in the rangi of 70 Wmv and speeds up to 12,000 rpm. The pump drivc nyb eurdt a4aserscint aif h requirement that no catastrophic damage be done to the Matia drive train in UK. event of acomsory failure. rte ffitar systwn; should conornst or a pump inlet icrwn to prevent .ngretiny n larar nnrtirLpcs anej, downstream of the pump, liner filtrartk. A primary disy-miablc fjilter element of required fineness in a housing with full flow bypass capacity. and posibly with bypass indica'zor, shouldf be w'ajvided. Replacema-nt C this filter element wvill be part oi the periodic; maintenance requirements. Additional filters may be required in the system to meet filtration requirements of fuall bypass flow, if stipulatcd by tht RFP or PIDS. In this case a secondary filkte sytcni will be installed in the bypass sytmn to assure continued clean oil delivery to the gears and bearings subsequent to complete clogging of the primary filter. The seondary filrtrtion requirements generally are less stringent than primary filtration. Filter elements of the order of 40 microns suitable for secndary filtratiouý can be of porous bronze, steel mesh. (x papar elca-smt types. The, bconze and stel flmws are cleanable ATIn reusabl while the paw element xznsally is 4-43
0.
t~~
IR REIE VI
'I
I WA
I
TEMPERALTUREG
I
JE~THEMA BYPASS~E
FLr
E A
THERLBALABYPASS
FIdr ". 01 ytm ceai
dispaiibl. The papr ekintrt, a-eued morei c
twod
ovjoaanc The p3-
rascnb eo-oaedan
ois mahita~xed with Ins. than 15 1%. p-cssutc drop wac ths; fi.r (Ref. 116). However, arnsolute filLuatiur of *~I5-Ir~ikian p&Aftnl with *9A, elfc.tuncy oi Fltering 5na~.con si~ze paililes or largoer h&. b%.~n shown to tw. coai~t4Cvte vtahaduquate (L.1 117). whaic finfr ftituoiv prommied filtci clomiat pwtulen,. that afkekd wevice intervahs *ud mlability.
x
444
boicoerstmnral)wicnitofn
an eter T1
startin,
oalirg aidroeftw i l coolist o cooerstm
dtuig~i requirements and pixocdurts arc prtsentcd in C..apter 8. AMCP 706-201, and are amplified in pur. 4--t.2. Forctd atr ztin be fron. shaft driven blowers or bkeec air turbiggc". 0-1 cuoole fen design~ procedui,; a~bo i. 4secribd ini Chap. -r 8. AMCP 7W2021Il. Cooke. k, :a1ion shoulf be cuiauabcnt with ap. p&mbie ballsauc thnca 1 survi 4bitiiy requirements.
C 7W202
C, Coolers integral with the gearbox or, if separate, surrounded by protective components or structure are possibilities. The use of either auxiliary systems or armor plating should be considered only as a last resort. The integral oil cooler has been shown to be effective (Ref. 116). It virtually eliminates the need for external plumbing and minimizes the ballistic threat !o the gearbox oil system, and the inherent protection of the surrounding airframe structure is enhanced by proximity of cooler to gearbox. A pressure bypass has been used to divert full oil flow to the gearbox oil system to circumvent oil flow to the cooler in case of ballistic strike on the cooler. The manifold is an oil distribution mechanism that normally houses the oil pressure regulator, temperature sensor, pressure sensor, and distribution passages. 3il is carried from the manifold through gearcase internal passages to oil jets for pressure lubrication of gears and bearings with direct impinging streams of high velocity oil. Internal passages also can be provided to direct oil to bearings encapsulated in housings and liners, Externally mounted oil system components such as pumps and filter housings often present sealing problems and service problems associated with the seals. Gaskets and O-rings normally are used for sealing between the mating parts. Components requiring frequent removal can more effectively be sealed with 0rings then gaskets. The compressed gasket material adheres to both the sealed surfaces, and mechanical removal of the gasket residue often is required. This becomes more difficult around studs: Each gasket application generally is unique and hence maintenance support requires stocking of unique parts, while 0rings are stocked for multiple applications and are supplied from a common stock. Gaskets possibly have a cost advantage by virtue of the elimination of the O-ring groove. Provisions for 0-rings also can result in slightly higher weight than for gaskets. Some system protection is provided by gaskets by their inherent ability to "blow out" in case of oversurges. Where close tolerances must be held between locations within the mating parts, the use of a gasket becomes impractical. The gasket material can compress and generally is not consistent from one gasket to another. In this case an O-ring should be used. c 4-4.1.3 Special Considerations S High flow oil systems may require multiple clemen^ pumps. Constriction free inlet design, high rotational speed, and high flow rate may not be attainable with a single element pump. Multiple element pumps (or more than one pump) also may be
Spressure
necessary in a dry sump design, when a scavenge pump is required. The scavenge pump extracts oil from the sump area and feeds it directly to the pressure pump or to an oil inlet sump for the pressure pump. Oil return requirements. sump capacity, and turn-around time must be compatible. A 20-gpm flow requirement with an 8-qt sump capacity results in complete turn-around of the oil ten times a minute. If the height of the gearbox is appreciable. with extensive baffling, there is a danger of pump cavitation and interrupted lubrication. Even without the danger of interrupted lubrication the oil has insufficient dwell time for deacration. Therefore, excessive foaming and inadequate lubrication or cooling are possible. Turnaround frequencies greater than 3.5-4 times per minute become questionable with respect !o proper deacration and attendant cooling characteristics. Adequate film thickness is difficult to achieve relative to surface finish in loaded contacts when low viscosity synthetic oils are used for gear and bearing lubrication. Boundary lubrication states, often characteristic with low viscosity oil, can still provide adequate wear life in gear teeth and bearings but the surface roughness must be low enough to prevent progressive metal-to-metal contact (see par. 41.2.1). Synthetic oil, especially MIL-L-23699, has a moisture absorption capability and its lubricating ability is diminished by moisture content. Hence, extreme care should be exercised in the design and location of gearbox vents to prevnt water ingestion. Areas where atmospheric air can impinge directly on shaft seals also should be avoided. Positioning rotating shields in front of shaft seats is a very effective ,means of preventing dust, dirt, and moisture-laden air from being ingested into the gearbox. Secondary effects of moisture absorption are internal corrosion. Synthetic lubricant that is contaminated with moisture becomes highly corrosive to the bare steel parts inside the transmission, with the lower roughness surface finishes being particularly suisceptible. It should be noted that once contaminated with water, MIL-L-23699 does not release that water when heated to normal operating temperature (>212°F). Therefore, both the poor lubricating quality and adverse corrosive tendency are present, and every effort should be made to prevent moisture absorption. The most consistent problem facing the designer of oil lubricated gearboxes is proper sealing. Leaking seals represent the single largest replacement item or cause for removal of gearboxes in the military helicopter (Ref. 33). Although it is infrequent that a seal leak rate is sufficient ior depletion of the gearbox 4-85
lubricating oil to occur in a single mission, that appearanc is presented nevertheless. The oil residue from a leaking shaft meal accumulated on the sur!uding components is so extensive that a minor eak manifests itself - a major problem. Certainly effective mad designs are laboriously, if ever, achieved. Carbon face and circumferential seals required for high-speed an,d high-temperature applications generally require an extensive test and development prolpam. Elastomaric shaft seals for lower spe applications are deskned more easily but successful sealing often is equally0difficult to attain, Investigations are being c€-ductmd continuously by and tm.-s to devlop a useal manufacture•rs versally acceptble and usersectov des. Based ae on versally acceptable and effective seall design. on the premise that no seal is completely effective, one design approach that can be taken to minimize the leakage problem is multiple seals. A shaft seal of conventional design, eithi elastomeric lip seal or carbon face smal, can be used in conjunction with other type of seals to affect seal staging. One suitable method is to use an inner lip seal with an outer labyrinth seal. The oil lubricates the lip seal, which assures adequate seal life, while the labyrinth provides secondary sealing from both directions. The shielding effect of the labyrinth precludes atmospheric debris that wouldbnomall arcceluere eastmomerc s haft we would normally accelerate dlastomer ades and shaft wear from collecting on the lip seal. A rotating slinger in close proximity to the housing on the outside will prodce affingand urter
ncrasetheseaing
effectiveness and seal life. Oil that weeps past the lip seal in normal operation can be removed through an overboard drain. veroardh d arains lp Research with various lip contact configurations for rotating shaft lip seals has shown promise during testing but no striking improvement has been observed in service. A ribbed lip was observed to produce a pumping action that prevented oil flow from the oil side of the test gearbox. Another lip design, a waved contact lip, produces a 3imilar wiping action and retains some lubricant on the seal-shaft contact that provides good sealing and coincident lubrication. A radially segmented carbon seal has been extensively tested and evaluated at NASA for high. This seal consists of several semispeed shaft sealing. rited ogeherand prig-ladc to circlarsegent circular segments fitted together and spring-loaded to
contact the shaft. In operation the seal lifts off slightly and virtually frictionless contact results. COOLING REQUIREMENTS Determination of the power loss in bearings and gears as described in pars. 4-2.2.1 and 4-2.2,2 provides the basis for determination of minimum heat rejection requirements. The gearbox frictional losses 4.4.2
"4-86
and windage loaes, having been determined, an estimate of the oil flow requirements can be made. 44.2.
Heat ExeP gh
Sizag
The maximum size renuired for a heat exchanger. or oil cooler, would be that size necessary to reject all the heat losf from the tranmnission. On the other end of the spectrum, considering forced air convection around the gearbox, no oil cooler may be required. This would occur if the surface area were sufficiently large, heat generation low, and internal oil flow distribution such that transfer of heat to the housing insiderate3 wallsfrom werealuminum adequate.and Characteristic heat transfer magnesium heat trans s fro in an d magnesu gearbox housings are in the range of 0.001 Btu/in.'man-F (Ref. 4). Hence, when the friction and windae loss has been determined and the surface area has been established, it easily can be decided whether an oil cooler will be necessary. Tail rotor drive gearboxes and accessory gearboxes generally fall into this category. However, in the interest of compact design it is rare that no cnoler is required for a main rotor gearbox. A general design requirement for the cooler is to reject 67% of the heat generated from the gearbox during critical Critical op rto.gn rly...sdoperating riconditions. g h vr td sg operation generally occurs during hover at design gross weight in hot-day conditions (35C. 4000 ft), when main isrotor power As is required and forced maximum air convection minimal. the gearbox forer-onvectio inimalAs th gearo x in design and material technology, larger size coolers will be required to reject the increased amount of heat that will result from higher specific gear and bearing loads and decreased wetted areas of housings. The physical size and configuration of the oil cooler, together with oil cnd air flow rates and pressure drops, can be determined with the help of the cooler manufacturer. The cooler core size and density are determined by the heat rejection requirements and the available airflowq. The procedural approach to cooler size determination consists of the following: following: 4-2..2) arbte ge ls (par. bye t. Determing 4-2.2.2). (par. area of losswetted by windage bearings, and teeth. 2. Determine effective external
exclsetoe areages ffective 2. e of appendaiges. exclusive is gearbox. Effective area 3. Apply heat transfer factor, 0.001 Btu/in.'-min4F for hot-day performance and power condition, and take algebraic difference between heat generated and heat transferred. 4. If heat generated exceeds heat transferred, then a cooler will be required to reject the excess generated heat.
S. Determine location for cooling fati an" based
upon the bystcm interfaces, i.e., available fan drive povwr; location of fan and ducting required, fan size
without bemult of prunesm lubricatio and =oolig systems. Much wock has been accoomp~lued in a, toblijagm an evalmUang design pwwsomu associated with anwuaocy operatka. i~e., failwink dasign (Ref. 35. 37. and 38). The baskc criterion tha has. hm atablished as sift coatinuatice of flight for a minimum of 30 min subsequent to total kw. of lubricant. As a minimum. the continuing flight ,Auft at the power level required to maintain dhe sped for maximum range at sea level standard conditions. Lou. of lubricant initially is synonymous with toss of cooling and is followed iminediattly by an incrase iv the coefficient of friction with attendant inacaise in heat gencratsio. due to the chanre to dry operation. As the primary heat trmnstcr medium of oil is lost. an immediate heat buildup occurs at the heat-generating points; and the secondary beAt transfar paths becormt paramount. If unstabilized heafiing is to be averted, the heat generating element
te airiflow coo!= resulting y5arudn limi atioas onifaeo hardware; and moe location, and sine limitations. The -nccessaiy calculations for a cooler and fan design are presented
hat fha medium rtdt atliiain throgighaac the secondary transferred utmiti maximum temperature: that is safe. Heat sources (V=a and bearings) must be designed to minimize
an required heat rqecton rate of the cooler, choos a fan "htwill im the airflw reqnwamats of tine ooe.C~oler specaicatkoa tw be met ame: * Rate oil fkow gpm Ib/mmi RAWU air Plow 0 Rated heat r4inction Swu/mmn * Oil iialsi teahareature OF * Oil Out- a temperature OFbe Psig * Oil inletpressure "*Oil Pressure drops Psi in. HC Air static premair. drop (Thmlew values will be establishe by the systan design and the heat rejectioni requirements.) Army number of actual ecombinations of airflow and cooler size can mies the establishe heat rejection requirements. Hlowvecr. the final chicke will be based
J
I441.2 wl
Casting F
*The
*
iitak
Sizing of the cooling fan can be accomplished by the method described in Chapter 8,AMCP 706-201. airflow requirements (volume, preusurc. and velocity) will be determined by the heat rejection required of the cooler, the oil flow rate, and cooler core parameters. Based on the required airflow, a ran that will interface with the available drive and space can be designed to nmeet the requiren.ents. Both axial and centrifugal flow fans arc uscd in cooler blowers. Choice of the type. of fan is dependent upon airflow volume and pressure requirements. The axial fan generally is used where: higher shaft speeds are avail, able, and pressure head at the cooler ishigh. The ce.-ntrifugal flow fan generally is used where high volumt flow at lower pressures is required. An adverse sideeffect possible with the axial flow fait is a high pitch nloise.
transferred away by the most efficient means availGears dasig~icd for fail-safe operation must have: sufficient clearance to prevent interference at the highei stabilized temperature. The clearancer necessary is de-termined by calculating the diffecrential expansiori bctween steel gear centers and the same distance in the housing, which is usually no-sierrous material. For instance, a gear set consisting of straight spurs operating at a center distance LCD of 6.0 in. and a normal operating temperature of 200 *F may attain a temperature of 900OF while the aluminum housing containing the gears (and bearings) only rises to 4001F. The rate of expansion of the steel St is 6.5 X 10-' in./in.-*F and the aluminum expansion rate Sf, is 12 X 10-6 in./in.-*F. The differential amounat of expansion would then be 4Ž.CD
=
(A TF~,8, - ATAI -L(Ti
4-4~.3 EMERGENCY LUBRICATION Design of a power transmission system to meet specific emergency operation rcquircinents entails a coruprehensivt evaluation of each and every dynamic component that ca-i influence the loss of drive continuity as a result of ii~terruption or loss of lubri-
)
cation. Redundancy of power paths, dormant ataxiliary lubricants, secondary cooiing systems, and specific design tolerances are considerations directly pertinent to emergency lubrication or operation
&fMIEJfi~
1%dmjt.
Ina MI.inIble.
.. -,6,
AA,) 1K 4(7'2 - TI)A, 6 AI 1
X( LCD
-1(9W0 - 200)(6.S)l0-1 -(400 - 200) (12)10-6) X6.0 - [ (700) (6.5 X 10-1) -(200) (12) (10-')) X 6.0 (4-6t') -1(45.5)10-' - (24)10-'] X 6.0
=0.013 in.
The: significance of the preceding calculation is that
~te gears expand in a radial directiorn toward each 4-8?
-~-7
et16r 10 Wediia u jgwace More than the hammeg mmpeade to sepwafe the Ipars. In the eamesk to Am m1r opwabatn at fth mumed coosu itio w~bow isetwamfe the teeth would have to be cut (osmoler rocit diameters) aad/or the outside it - dKMWAd equivalimt to spraeading the gar -oiie by a toal of OA1l3 im. The amm typ ofcalcamlatios. ca... ot made for ball and roilsr bmains ca a radial clerance book and for deplex bail baimfsa. cons~hisig cootat a"Q.. on a radial aasd Wms. Raia grwt diromu in bearins eand gemr and the effects of dry rumuia~g an presustod in Ref. 38. Opeimam deuign far rani-as Ihictiom loss in Vars and bearings is covered in per. 4-21. To operat. a gma or beatin at temperatures or 9W F id above, a i aacsr that tb copnn be fabricated from material(s) that exhibit a "seMoMbl tolorance to high Wimperature. Materials N)and ANS 649 (MykAa AMS 647 (Nitrali;.. 9O) arn weil suited to the purpose. AMS 6M9 exMAUexalestbo brdm carctritic ad as Proves to be one of the most fatigue resistant beawaug mawariols available. AMS 6475 is a precipitation
-Ws
~
*rolling
.~i.L~h. a.
~
~4.
which also exhibits high hot hardness charom-e Ad%,a The more common geu mand beana& materials, AMS 426 and AMS 6444. respectively, do no have high hot hardness tLarlicte istes. but they cani withstand smoderate lod ror a short period. Fail-safe operatim for V wincabe obtained using AMS 265and ANS 6W4 parn and bearings, but the applications amalimited to moderate power levels wnd speeds. design of failBearing cages are also critical to tA~k safe brings ilrounz and plastic materials i-re not acceptable for fail-safe operation. The characteristic aildure mode for a bearnog with a bronze cage is mechanical plating of the broaze onto the rolling dclments with immediate loss of running clearanr- and tenperatu instability, followed by seizurc. Plastics such as nylon, Teflon (teirafluorotthylene) and fiberglass offer littlc resistance to failure at elevated tmutperaturcs. Carbon graphite is an excellent cage material for dry operation, but its tens~ile strength is too low for normal use. Mitnufacturing problems and scrap rates are significant; carbon cages can be armoend with steel reinforcing rings and side plate%but the cost is quite high. The Allost adaptable cage material at the present time appears to be mnild steel with silver-plated pockets. Dry friction between the elements and the silver-plated cagc is moderate, and the stoel retains adequate: strength for "ti application. Clarancez ar,-necessarily a very important part of the cage desiign. Outer-land-riding 4-38
COWm offa tOw risk of entrapping slag-type debriss betwasa ilhc cage and the outer land. with fracture cc soma po~ible, while inner- lmnd-riaing cages risk loss of dleautmm dues to thernal iflerviatiaks between the cage and tlw saner inag of the bearin. The most difficul cope deep issen Wn high speed bearnogs. It is desirable especially in roller bearings. to provide imaer4and-riding cages with inne ring thrughhabrication to make maximuam advantage of the traction force vectors. Also, it is desirable to minimize guui P ng-d-to- cage dlearamc from a dynamic balance stiandpoint. Therefore. if normal high wooed deigin pareameters are followed, risk of seizure; or *bun-ut*O of the beariSg Weream5 for di y opera
-
&M.
Seveal means of augmenting lubricatioti or supplying lubricant after loss of the primary oil system, that may be developed are: I. Inclusion inside rotating shafts of high melting point lubricant tha melts mand flows into bearings and on6o gears aftes dry running commen~ces 2. Providing oil traps with metering holes 3. Wicking oil into bearings from oil absorbing maiterials Encansulatina lubricant in containers with heat acuvaelo drain pangs 5. Prokt" awuxliary (idler) gears of oil absorbing or dry lubricant material to nmes with power
4-
j
ACCLSSORIIES
4-4.1 PAD LOCATION AND DESIGN CRIrMIA On small helicopters the atccessories may consist only of an oil pump. hydraulic pump. tach generator, and Cooler fan. A simple cc-axiai ar-rangement of oil pump, tach generator, and hydraulic pump a"on the OH-58A may be tlhe most effective means of arranging an accssory drive. The accessories arm driven by a concentrated contact spiral bevel pinion powered by the input bevel gear. On medium weight helicopters, where System TCdundaitcy may be required, multiple: accessory pads usually can be provided on the main gearbox. Hydraulic pumps for primary control actuation must ac located at widely displaced locations to thwart loss of both systtms to a single small arms bullet . The size of the main gearbox normally will be adequate to allow such displacement while etill providing pads for generators, tach drive, etc. Accessibility for maintenance must still be a prime critericn for location. On. large helicopters the most effect~ve means of providing accessory drives normally is frown a gear-
j
AMCP 706-202 box remote from the main rotor gearbox. Multiple with redundancy become imperative, and the complexiay and power required fok- ground checkout estabiAhes theu sed for an auxiliary power unit (APU). With multiple drive pads and high continuous power requirtment tho remote accessory drive gearbox must have a recirculating oil system,. coplete with oil pump and filter. For emergency lbrication considerations, tle gearbox must be selfcontained to prevent oil depletion from the main tramnsibsa in the event of the occessory gearbox being hit by small arms. The location for the meomozy gearbox must not introduce unacceptable noie kvcls in crew compartments.
rsyhtems
Spower
)
482 ACCESSORY DRIVE DESIGN REQUIREMENTS ancofgrtofc-lo.Je Accussory gearbox desin adcniuainfc tors must be compatible with the main gearbox takeoff. airframe and cowling, work platform provisions& CGI. and minimization of gear-induced ni Particul~ar drive pad power requirements arc determined by the accessory (hydraulic pump, mneratort alternator; eittA and dhw nrn.er Wq4 AND, or QAD pad must be provided to ameet the continuous power rating and seizure torque level. The ocavuory driiee shaft normally isprovided with a sersection that must fail in the evv:nt of seizure of Ohe accessory rahrthan pemtdamage tothe c192 onssory gearbox. Coincidentally the accessory drive gearbox must be provided with a connecting drive buha&t system that will isolate elfectu of accessory gecarbox seizure from the main rotor gearbox. Multiple dutc arrnge a~ mreqie to poieisolation of the APU during normal helicopter operation and _.u....
=--
e~rl. &m1.001N
rotor gearbox duiing APIJ drive (ground checkout, c.). APU shaft mounted centrifugal clutches am welsuited to the former application and one-way qpfaf ltce are wdl suited to the latter. Functionally, the APRJ must power the acccusoy gearbox by driving through the APU input clutch while the main gearbox drive is disengaged by neans of the APU is sbut down and disengaged while the main rotor gearbox drives into the accessory gearbox through the one-way clutch. Additional clutches may berequired to limit the number of accessory drives tht operate during ground operation.
gearbox tocation. The hydraulic pumps arc especially severe noise geenerators and close proximity to a crew compartment can cause intolerable high-pitch sound levels. Elastomeric mounts can be an cffcctivc noise isolation means. APU exhaust ducting must be adequate to pievent noxious gas and heat from invading the ptrsonncl compartments. As with other gearboxes accessibility must be provided ti., oil lcvcl indicators for preflight maintenane. One man should be able to change accessories without assistance. REFERENCES 1. Darle W. Dudley. 77#e Evouioa ofthe Gear AAn.
circa 1966, AGMA 990.14. 50th Annual MEeeting 9. c,f the American Gear Manstfactwrers .lssociaa.Jn 96 2. D. W. Dudley, Ed.. Gear HamdbAoo. McGrawHill Book Co., NY, 1962, Chapter 5, p. 20. 3. Mbd. Chapter 14, p. 20. 4. Insallation of Hlgrh-Adoction-Ratio Transsmu-
sio.. in the UH-l Helicopter. USAAVLADS TR -. &
fe
S. F. A. Thoma, Written commentary on paper, Thkenwwl Bdhawlo, of High-Speed Gears by Luigi Martinaglia. ASidE Symnpomium an Trensmitsions A Gearn, San Francisco, CA., October 6. Arvid Palmgrca. B.li and Roller Bearing Enginee'ring, S. V. Burbank A Co., Inc., Philadelphia. PA.. 1954. pp. 36-41. 7. Donald F. Wilcox. and E. R. Booser. Bearing Design aid Application. McGraw-Hill Book Co.. Inc., NY, 1957. p. t3. -. Tedri..k A. Harris Rolling Bearing Analysis. John Wiley A Sons. Inc.. New York. 1966. Chapter 14 and 15. 9. A. B. Jones. "Rai Motion and Sliding in Bill Bearing&"~. ASP4E Journal of Basic Engineering. 1-12 (March 1959). 10. P. D. Waehler. The Ralway Conical Bearing. RE-0012-01. Rollway Sc~aring Co., Inc.. 12. Predlctod Characte'ristics of an Optimized SeriesHyl'rid Conical Hiydrostatic Ba11 Bearing.
NASA TN D-6607. Decmber 1971. 13. "Lubrication and Wcar". Lubrication. Texaco, Inc.. NY 31. No. 6 (1965).
ý"_ASPEIALREQUREMNTS14. R. S.Fein and K. L. Kreuz. Disnasion on BowmSPECILREUIREENTSdary Lubrication, NASA Symposium team
hemutbcosdrdin thcoieof
TX, November 1967. pp. 6.2.1-6.2.25.
15. D). Dowscn, EJamswyd'odynamlc Lubrication. NASA SP-237, Symposium on Inmerdis.-plinary Approach to thge Lubrictotion of Concentrated Contoas. Troy, NY, July 1969. 16. A. Ott, Elast ndrodytwam'ic Lubrication of Involute Gears. ASME Paper 72-PTG-34, Mechanisms Contference A International Sympaslum on Geafin aod Transmissions, San Francisco. CA. October 1972. 17. E. 1. Radrimovsky. A. Mvirarcti. and W. E. Broom. Instantaneous Efficiency and Citfficient of Friction of an Involute Gear Drive. ASM F Paper 72-PTCF-13, Mechanisms Conference A Interniational Symposium on Gearing and Transmissions. San Francisco, CA, October 1972.
28. C. W. Bowen, Analysis of Transmission Failure Mc~es. SAE Paper 710454. Atlanta, GA, May 10-13. 1971. 29. Designt Contact Stress Limit Rleconmmendations for Aerospace Gearing. AGMA Specification 411.02, Septenmber 1966. 30. Bowen, op. cit., p, 5. 31. 3. A. '3e~n and J. H. Tinggold, Results of the Reliability and Maintainability Dem~onstration~ of the OHl-58A Light Observation Helicopter. A HS 28th National Forum. Washington, D-C, May 1972. 32. CH-54 Reliability and Maintainability, U.A.C. Report 64276, USAAVSCOM Contract No. FDAAJOl.68C05l2(3I)P008. Novcrn~bcr 1972. 33. Identification and Analysis of Armyv Helicopter
18. L. U. Tso and R. W. Prowcll, A Study Gf Fri c-
Reliability and Maintainability Pro bi ems and
lion Loss for Spur Gear Teeth, ASM E No. 61 WA-85, October 1962. Dudley. op. cit., Chapter 14, p. 5. H. E. Staph, P. M. Ku. and H. J. Caper, Effect of Surface Roughness and Surface Texture on Scq~fflng. ASMF-AGMA-IFTMM Symposium on Gearing and Transmissions, San Fraricisco, CA.. OactoblCr 1972. Heat Generated in High Power Reduction Gearing. Report No. PWA-3718 prcpar-d under Contrart No. N00019-68-C.0422, Naval Air Systems Command, June 1969. Analysis of Noise Generated by' UH-l Helicopter Transmission. USAAVLABS TR 6841, June 1968. fnisms, Program for Helicopter Gearbox Noise Prediction and Reduction. IJSAAVLABS TR 70-12. March 1970. Thomas Chiane and R. H. Hadglev. Reduction of Vibration and Noise Generated by Planetary Ring Gears in Helicopter Aircraft Transmissions. ASME Paper 72-PTG-l I, Mechanisms Conference A International S~ymposium on Gearing and Transmission~s. San Francisco. CA, October 1972. E. 1. Radzimovsky and W. E. Broom, Efficiency of Gear Transmissions with Flexibility Connected Grars Subjected to Axial Vibrations. ASME Paper 72-PTG-I0, Mechanisms Conference & International Symposium on Gearing and Transmissionts. San Ffancisco. CA. October 1972. R. G. Schlegel, R. J. King, and H. R. Mull, GCear Noise", Machine Design. February 1964, Study of Helicopter Transmission System Developmetnt Testing. USAAVSCOM TR 69-3. June 1968.
Deficiencies, USAAMRDL TR 72-1 IA. Vol. 1. April 1972. 34. "Automatic Control for Hydraulic Systems", U. S. Patent No. 3,474,819, October 1969. 35. Solid Lubricants for Helicopter Tail Rotor Gearboxes. Final Report, Contract DAA.J02. b'GW058, US Army Aviation Material Litbora-
19. 20.
21.
22. 23. 24.
25.
26. 27.
~4-90
toricSi.
36. Vulnerability Study of_ the UH-) Helieopter Power Train System to Small Annms Fire. (U) BLR Memorandum Report No. 1821 (C) February 1967. 37. Gre~ase Development and Evaiuati..n for Helicopter Transmissions and Servo MechaUSAF Tcchnical Report AFML-TR-68338, Parts 1 and ii, November 1970. 38. Fail-Safe Bearing and Gear Lubrication System. Final Report, Contract Now-65-0592-ci. Bureau of Naval Weapons. Washington. DC. March 1969. Edac fCruie 39. C. W. Bowen, Pilting duacofCrrie Spur Gears in Synthetic Lubricants, AGMA Technical Meeting. Chicago, I1- November 1967. 40. M. A. H. Howes, Rielationship of Lubrication and Fatigue in Concentrated Contacts. ASME Report No. IITRI-B8l 35-S. NY. D~ccembcr 1971. 41, C. W. Buweit, Strength of Gears - New~ Materials Investigation. AGMA Aerospace Gearing Committee Meeting. Mil~aakcc. W1. September 197042. F. G. Rounds. "Some Effects of Additivecs on Rolling Contact Fatigue". ASUL Trans~actions 10, 243-255 (1%?7). 43. G. 1. Giaham, Combat Optriaftnal Hlight Profties an the UN-IC. AN-IG. wod (li1-Ill
F' ____
_
____ ____
__7
____
___
AMCP 70&-202 Helicopters. AHS. 26th National Forunt. Wash-
ington, DC. June 1970.
the Effect onI Their Load Carrying Capacity. N. Report No. 102 (Part No. 3), Department of Scientific and Industrial Rcscurch, Sponsored Research (Germany), 1947.
44. Trade-off Study for Exteinded-Life Helicopter Transmijslons. USAAVLABS TR-72-40, November 1972. 63. Strength of Spur, Helical, Herringbone and Bevel Gear Teeth, AGMA Information Shoet 225.01, 45. Dudley. op. cit., Chapter 14, p. 56. D~ecember 1967. 46. Mode of Failure Investigations of Helicopter 64. "Bending Stresses in Bevel Gear Teeth". Transmissions, USAAVLABS TR 70-66, November 1972. Gleason Works Gear Engineering Standard; "iribos", Tribology Abstracts. The British Rochester, NY, 1965. h-ydromechanics Research Association, Cran65. E. J.Wellauer and A. Seirig, "Bending Strength field, Bedford England, published monthly. of Gear Teeth by Cantilever-Plate Theory" 48. J. 0. Smith and C. K. Liv, "Stresses Due to Journal of Engineering for Industry, 82, No.3 (August 1960). Tangenwial and Normal Loads on Elastic Solids With Application to Some Contact Stress 66. W.L. Mclntire. et. al., Bending Strength of Spur Problems", Journal of Applied Mechanics, £0, 2adHelical Gear Teeth, AGMA Paper 229.21 (June 1953). October 1967. Srnt 49. B. W. Kelley, "The Importance of Surface. 67. W. Coleman, A New Perspective on the Srnt Temiparature to Surface Damage", Handbook of of Bevel Geer Teeth. AGMA Paper 229.13, OcMtechanical Wear. University of Michigan tober 1969. Press, I1961. 68. Evaluation of Advanced Gear Materials for Gear-. 50. B. A. Shotter, A Newi Approach to Gear Tooth boxes and Transmissions. Firial Report. ConRoot Stresses. ASME Paper 72-PTG-42, Octract N00156-69-C-1965, Department of the tober 1972. Navy, NAPTC (AED), Philadelphia, PA, 51S.A. Y. Attia, Bending Moment Distribution on Septemiber 1961. IF Gear Tteeth 01 Llrcidnkr-arc Frojiles. ASML 0Y. J. Bi. Seabrook and Di. W. Liuficy, JAesuirs OJ a Paper 72-PTG-46, October 1972. Flftezn Year Program of Flexural Fatigue 52. Dudley, op. cit., Chapter 14, p. 43. Testing of Gear Teeth, ASME Paper 63-WAJ. R. Miller, Fillet and Root Design Con199. November 1963. siderations. AGMA Paper 109.26, February 70. E. J. Ripling and J. E. O'Donnell, How Fracture 1971. Mechanics Can Help the Designer, SAE Paper .54. Advancemeni of Spur Gear Design Technology. 710153, January 1971. USAAVLABS TR 66-85, December 1966. 71, H. J. Carper, P. M. Ku, and E. L. Anderson, 55. Advancement of Helical Gear Design TechEffect ofSome Material and Operating Variables nolagy. IUSAAVLABS TR 68-47, July 1968. on ScuffIng, ASME-AGMA-!FTMM Sym56. Advancen:ent of Straight and Spiral Bevel Gear poslum on Gearing and Transmissions, San FranTechnology, UISAAVLAES TR 69-75, October cisco, CA, October 1972. 1969. 72. H4. Block, Les tempemnuires des surface dans des 57. A. J, ~.,nianski, Gear Design. SAE 68038 1, Fifth condition de graissage so=s presslon extreme. Southern New En~glitid Seminar. April 1968. Cougr. Modial Petrole, 2mc Congr., Vol. 3. 58. A. 1. Tucker, Dynamic Loads on Gear Teeth, Paris, 1937.
I47. I F53.
Design Applications. ASME, Design Engineering Conference and Show. NY, April 1971. 59. A. Seirig and D. R. Houscr, Evaluation of Dynamic Factors for Sp,'r and Helical Gears.
ASME Paper 69-WA/DE-4. November 1969. 60. D. R. Houser and A. Seirif, An Experimental Investiration of Dynamic Factors in Spur and
Helical Gears, ASME Paper 69-WA/DE-5, November 1969. \ 6,'. A. Y. Attia, Defection of Spur Gear Teeth Cut ] in Thin Rims, ASME Paper 63-WA-14, 6.November 1963. C. Wcber, The Deformation of J.oaded sgearqr and
73. Gear Scoring Design Guide for A erospace Spur awed Helical Power Gears, AGMA Information
Sheet 217.01, October 1965. 74. V. N. Borsofi', "On the Mechanism of Gear Lubrication", ASME -journal of Basic iinginecring, BID (1959). 75. P. N. Ku and B. B. Baber, "The Effect of Lubricants on Gear Tooth Scuffing", ASLE Transactions, 2 (1959). 76. B. W. Kelley and A. J. Lcmanaki, Lubrication of 1n;,oiurr'Gearing, Coqference on Lubrication and Wear, Institute of Afechanicat Engineering, Lorn-
don, September 1967. 4-91
AMCP 70W202 77. D. W. Dudley, "Modification of Gear Tooth", Product Engineering. September 1949. 78. H. Walker, "Gear Tooth Deflection and Profilc Modifications", The .nglneer, London, 3 Parts, October 14, 1938/'October 21, 1938/August 16, 1940. 79. R. Pedersen and S. Rice, "Case Crushing of Carburized and Hardened Gears", SAE Transactions. 370-380 (1961). 80. Surface Durability (Pitting) of Spur, Helicai, Herringbone. ard Sew/ Gear Teeth, AGMA Information Sheet 215.01, September 1966. 81. How To Test Bevel Gears, Gleason Works, Rochester, NY, 1955. 82. C. W. Bowen, Helicopter Transmission Design, presented to Texas SAE Region, 1961. 83. A. B. Jones, "Analysis of Stresses and Diflections", Ncw Departure Engineering Data, 1, 161 (1946). 84. T. Harris, An Analytical Method to Predict Skidding in High Sperd Roller Bearings, ASLE Paper 65-1-C-14, Park Ridge. IL, October 1965. 85. "Anti-Skid Bearing", U. S. Patent No. 3,410,618, November 1968. 86. G. Lu~ndbe'rg and A. Palmaren. Dynamic Capacity of Rolling Bearings, Acta Polytechnica, Stockholm, Sweden, 1947. 87. G. Lundberg and A. Palmgreri, Dynamic Capacity of Rolling Bearings. Acta Polytechnica, Stockholm, Sweden, 0952. 88. W. Weibull, "A Statistical Theory of the Strength of Materials", Proceedings of the Royal Swedish Institute for Engineering Research, 151 (1939). 89. Bamberger, c', al., "Life Adjustment Factors for Hall and Roller Bearings", Engineering Design J-
AC'&AV %usuaur MOML, S'
NY,
N~
r,,~
Snt,.mber -
197!.DfetoDanss1adPonssi
90. A. B. Jones, "A Ger'eral Theory for Elastically Constrained Ball and Radial Roller Bearings under Arbitrary Load and Speed Conditions". A SME Journal of Basic Engineering, 309-320 (June 1960). 91. A. B. Jo'ies and T. A. Harris, "Analysis of a Roll'ng Element Idler Gear Bearing Having a Dtiormnable Outer Race Structure", ASME Journal of Basic Engineering. 273-277 (June 1963). 92. Charles Wilson, Curvic Couplii g Design. Gleason Works Design Guide, Rochester, NY. February 1964, 93. John Kayser and Wilson Groves, A New Concept in Dpive-.'ine Slip Splln-'s. SAE paper 680118, January 19t8. 94. E. A Ferris, A utomotive Sprag Clutches -
4-92
Design and Application, SAE paper 208A, January 1901. 95. Sprag Overriding A ircrafi Clutch. U SAA MR DL TR 72-49, July 1972. 96. Aircraft Clutch Assemblies, Ramp Roller. USAAMRDL TR 72-31, July 1972. 97. Spring Clutch Applications. Enginecrisig Report No. PD-462A, Curtis5-Wright Corp., Caldwell, NJ, February 1964. 98. Spring Overriding Aircraft Clutch. USAAMRDL TR 72-17, May 1973. 99. "Hydraulic Brake", US Patent No. 3,228, 195, January 1964. 100. Crash Survival Design Guide, Revised, USAAVLABS TR 71-22, October 1971101. "Surface Scaling"', This is Magnesium, 16, Heathcote and Coleman, Birmingham, England, August 1968. 102. J. H. Hull and S. J. Erwin, The Effect of Mechanical Deformation on the Tensile Properties and Residual Stresses in Aluminum Forgings. ASME Paper W72-53.1, Western Metal and Tool Exposition and Conference, Los Angeles. CA, March 1972. 103. Danle W. Dudley, Successes and Failures i.n Spac. Gearing. S:AL-AbME Paper No. 671lB. presented at the Air Transport and Space Meeting. NY, April 1964. 104. Wayne L. Mclntire, How to Reduice Gear Vibra. tion Failures, AGMA Paper, presented ai the AGMA Aerospace Gearing Technical Cornmit tee Meeting, Orlando, FL. February 1964. 101. An Investigation of Helicopter Noise Reduction by Vibration Absorbers and Dampirg. USAAMRDL TR 72-34, August 1972. 106. Rudolph H-ohenberg, Characterization of alr elcin igoiadPonssi alr Prediction, Proceedings of the Tenth Meeting of the Mechanical Failures Prevention Croup, Ofrice of Naval Research, January 1970. 107. Donald Davis, Time Series Analysis Techniques. Proceedings of the Tenth Meeting of the Mechanical Failure Prevention Group, Office of Naval Resea.ch, January 1970. lO8, J. M. Vanct, Influence of Coupling Properties on the Dynamics of High Speed Powei- Transmission Shafts, ASME no 72-PTG-36, Mechanisms Conference and International Symposium on, Gearing and Transmissions, San Francisco. CA, October 1972. 109, Flighs Test Evaluation of a Supercritical-Speed Shaft, USAAMRDL TR 70-50, September 1970. 110. Design Criteria for High-Speed Power Transmis-
l
•.
-..-
.)
1.. I11. 112.
113. 114.
1..
. ...
sion Shafts. ASD-TDR-62-128, AFAPL, Wright-Patterson A: Force Base, OH, December 1964. The Bossier Coupling, NASA CR 1241, National Asronautic, and Spary Adninistr1tion, Washington, DC, January 1969. The Bossier Coupling Experimental flight Test. Finai Report, Contract No. 0156-69-C-1316, Department of the Navy, Naval Air Propulsion Test Center, Washington, DC, March 1972. C. W. Bowen, Gear Couplings, AGMA Paper, Aerospace Gear Committee Meeting. Seattle, WA, September 1962. Comparative Lubrication Studies ojf 0.1-58A Tail Rotor Draveshaft Bearings, NASA TMX68118, NASA Technical Memorandum, Cleveland, OH, July 1972.
-.
* -
--
,, ,
7-. .
AMCP 706-202 115. An Extreme Pressure, Anti-wear Grease for Transmission Lubrication. AFML-TR-72-282, USAF Technical Report, December 1972. 116. R. Cooper, Development of a Three Micron Absolute Main Oil Filter For the T-53 Gas Turbine, ASME Joint Fluidi Engineering, Heat Transfer. and Lubrication Coqference, NY, September a970. 117. OH-6A Product Improvement Program Upgrade Transmission to a Longer Life Configuration, Final Report, Contract DAAJOI-68C-I 123, p. 20, US Army, AVSCOM, St. Louis, MO, May 1973. 118. Investigationof an ExperimentalAnnular-Shaped Integrated Trarsmission Oil Cooler Design. USAAVLABS TR 70"4, September 1970.
4-93
CCHAPTEP 5
ROTOR AND PROPELLER SUBSYSTEM DESIGN 5-0 A a a,
LIST OF SYMBOLS
b b,
- propcller inflow angle, de3 - spred of sound, fps - cofficient which is derendent upon mass and stiffness distribution and has a differect value for each mode of vibration, dimensionless = tip !oss factoi, dimensionless - blade loading, lb/ft' - number of blades - blade semichord, ft
C
- empirical constant, dimentionless
CD
- mean rotor blade profile drag coefflcient, dimensionless
CL
=
f 8L
4 KE NCR L, I M
M Md, ,
less
M, MR
- mass per unit length of the beam, slug/in. - mass of spanwise increment at outboard
m,
- mass of spanwise increment at inboard end
end of blade (N), slug
sionless %P
mil
c
cd
c c,, i
of blade (c). slug
- coefficient of pitching moment, dimensionless o .. r.Ip
, diiac
n`1 n
-
-
distance fro:n beam neutral axis to outer
fiber, in. - airfoil section drag coefficient, dimensionless = airfoil section lift coefficient, dimensionless = = maximum section lift coefficient, dimensionless
load factor, dimtnisionless the number of vibratory stress cycles aca2nariruiard vI.rcc e nvr f
,-siocm,,oed
particular operating condition
thrust coefficirnt, dimensionless
CT I,
,
coefficient of lift, dimensionless
- nrean meL rotor blade lift coefficient, dimenCM
coefficient dependent upon mass distribution and the mode of vibration, dimensionless - kinetic energy, ft-lb = rotational kinetic energy, ft-lb wing lift, lb = length, in. - bending moment, in.-lb or ft-lb Mach iumber, dimcnsionl=es = advancing tip Mach number, dimension-
H
P P Q, Qp Q,,r
q
tail rotor rotational speed, rev/sec
gust load factor, dimensioniess = actual powet required, hp pressure, psi = engine torque, units as required -, propeller torque, lb-ft = main rotor torque, lb-ft 2 - dynamic pressure, lb/ft
D
- propeller diameter, ft
R
= propeller tip radius, ft
E
- modulus of elasticity; psi
R
=
4,%,
=
.F
El e F g HP0
1, Ie
i,,,5K
excitation factor, dimensionless
- stiffness, lb-in) - location of flappiig hinge from the center of rotation, in. - force, lb - acceleration due to gravity, ftisec2 - profile power requiid, hp - mass moment of inertia, slug-ft2 = momcnt of inertia, in.' - polar moment of inertia (per blade for a tail rotor), slug-fV - propeller mass moment of inertia, slug-ft2 - mass moment slug-ft o helicopter yaw mass moment of inertia, slug-ft' - ratio of total tail rotor thrust to net tail rotor thrust, dimensionless -T - notch factor, dimensionless - gust alleviation factor, dimensionless
1R,
RN
r r S SFP. .S/A T AT T/A T
T,, T,, r V
rotor radius, units as required main . L.d adi.,. f, = tail rotor radius, ft - outside blade radius, in. = radius, ft radius of curvature, in. Laplace operator, see-' = stall flutter parameter, dimensionless = ratic of blocked disk area to total disk area, dimensionless - thrust, lb - change in thrus', lb = tail rotor disk or thrust loading, psf - tail rotor th, ust required to compensate for main rotor torque, lb total tail rotor thrust required. lb total tail rotor thrust minus the fin force, lb propeller axis downtilt from wiog zero-liftline, deg - true airspeed, kt robo
5-1
AMCP 706202 V
V. W W W
w X X"
x Y
• 6, 6
S-
average velocity of contacting surfaces, fpm airspeed, kt indicated f = induced velocity, fps vertioal airspeed, fps S art uste weight, lb - aft adjustable weight, lb W/ forward adjustable weight. lb proprller weight, lb maximum allowable weight for abrasion strip. lb sr minimum allowable weight for abrasiob strip. lb -
=disk loading, b/ft2 = distance bet~wen center of main rotor and tail rotor-antitorque moment arm, ft = dynamic axis, in. A clearance between main rotor and tail = rotor bladc tips, ft = chordwise distance from blade leading edge to centroid of mass increment. in. =spanwisc distance from flapping hinge to centrois of mass increment, in. = rntroi S onlfmass incrment, , in.that rotor blade o g n , dvice = propnler blade iangle, deg = Taprp (Cblade =Tt''( ICL) = pitch-flap coupling angle, positive if pitch is decreased when the blade flaps up, deg = rotor blade angle, de8 advance ratio, dimensionless coefficient of friction, dimensionless rotor mass ratio = air density, slug/ft1 = standard deviation, defined P.s the rootof the dviations b.S....... points and the mean data tween individual = rotor solidity, dimensionless =
p Ir
$T
=
4' 4' 4,
9 S;, R11
blade bending stress, psi
= rotor blade solidity, dimensionless propeller inflow angle, deg = yaw rate, rad/sec 2 - yaw acceleration, rad/sec
= rotor angular velocity, rad/sec = tail rotor angular velocity, rad/sec = rotor tip speed, fps -
precession velocity, rad/sec
- propeller speed, rad/sec = natural torsional frequency, rad/sec = natural frequency of a rotating beam, rad/scc
5-1
INTRODUCTION
In general, all rotors and propellers arc mcchanical devices used to produce thrust by accelerating 3 fluid mass. They range in sophistication from simple two-bladed, fixed-pitch configurations to coaxial counterrotation systzms with individual rotor colicctive and cyclic pitch control The analytical techniques for all types arc very similar. However, there are minor variations in the definition of rotor-propoller nondimensional parameters whi,:h prove to be unimportant once it is realized that data can be transposed readily from one format to another. The overall performance of a rotor or propeller may be described by its tip speed, airfoil characterisWics, solidity ratio, and disk loading. Rotational inertia also is important to rotor design because it affects helicopter autorotational performance. Based upon selected vialues for these parameters, the detail design of the rotor is largely a task of optimizing the configuration in terms of the number of blades, flapping and inplane freedoms, dynamic response to externally applied cyclic forces, and the assurance the hardware can be built with a fatigue or serlife compatible with the design requirements.
The paragraph addressing "esign parameters reviews those preliminary design factors which will be converted to useful hardware in the design of the convertemn rotor The system. paragraph on rotor system kinematics dis. cubses the blade motions to be accommodated in the detail design; in particular, the flapping, leading, and blade-feathering motions. Typical rotor systemi cccommodat, these mctions by means of teetering, fully articulated, or hingeless hubs. The paragraph also describes a number of methods that provide for both cyclic and collective feathering of individual blades. The paragraph on rotor system dynamics addresses
the internal stiffness and mass distributions of the rotor blades, and the relative effects of these factors on aeroclastic stability, vibration response, flutter, ground resonance, and other phenomena related to system damping and periodic forcing functions. Also
covered in this paragraph are rotor responses to such transient excitations as gasts and acoustic loadings. The discussion of blade retentions include. the various means of attaching the blades to the rotor
hub. Among these are elastomeric bearings, tensiontorsion straps, and antifriction bearings. Also described are auxiFaiy devices used at the hub to alleviate blade forces associated with blade pitch, and the lag hinge dampers used to dissipate the excess energy of the inplane motion of the blades. Blade-folding provisions, both manual and powered, arc discussed as well.
5-2
The paragraph on rotor blades discusses trade-offs in blade geometry, such as airfoi! section and root-totip taper and twist, and their relationphip to the corresponding parametric analyses discussed in AMCP 706-201. Design considerations that provide for manufacturing simplicity, inservice adjustments of blade balance and track, and the blade materials and joining needed to position masses and Sstiffnessestechniques properly, are addressed. Also discussed are rotor fatigue Th system paragraph on lives. propellers
3. Radar cross section 4. Damage tolerance against a. Striking a solid object such as tree limb b. Being struck by weapon fire, either solid or HE 5. Repairability 6. Fatigue life 7. Weight 8. cost
deals generally with
Specific the performance parameters probably willvalues haveorbeen selected during preliminary
the design requirements for propellers and develops design considerations in the same manner as do prior paragraphs for roturs. The paragraph on antitorque rotors reviews the knowledge gained in recent years concerning the desirable direction of rotation, the flapping freedom required, the merits of pusher versus tractor configurations, etc. The advent of "flat-rated" enginetransmission systems with high-altitude capability has placed additional demands on tail rotor control power. Additionally, the airspeeds encountered in normal operation have increased markedly, creating adverse environmental conditions for tail rotors. Th.. , rn niher et problems, are discussed in light of the lastest knowledge.
design. CornpliaricL with the operational criteria is dependent largely upon the materials and method of manufactume, which will be selected during detail design. The design problem initially is broken down into the requirements for hover, high-speed level flight, and high-speed maneuvering and each is discussed independently. The total problem then is considered and some approaches are offered.
5-2
DESIGN PARAMETERS
..
.
5-2.1 HOVER Selection of the optimum hovering rotor involves all the performance related parameters listed previously. with the exception of advance ratio. Hover power is divided into "induced power" (thi chs.-geable to providing lift) and "profile power" ithat chargeable to blade profile drag).
The selection of rotor parameters is quite complex, as each major variable interrelate-s with all other vari-
5-2..l
rotor performance arc outlined in Chapter 3, AMCP 706-201. Included is a discussion of the tyne of parametric analysis required to optimize a rotor fj' * given group of perfoimance requirements. The discussion herein supplements that description of preliminary design procedures, with emphasis upon the considerations pertinent to the detail design phase. The parati etots that are considered in connection with rot,.r performance include: 1. Uisk loading 2. Blade loading 3. Blade tip Mach number and advance ratio 4. Number of blades 5. Blade twist 6. Airfoil section(s). For an Army helicopter that will be requirel to operate in the nap-of-the-earth and in combat, complianc only with specified performance requirements will not produce an acceptable design. Additional design criteria that may or may not be defined quan. titatively for a particular helicopter rotor include: I. Maneuverability 2. Noise
loading is described in Chapter 3, AMCP 706-201. And:disk more extensiv, discussion can be found in Ref. I. Disk loading frequently is determined by factors other than performance. Fo" example, a requirement , for air transportability may dictate a fuselage length ih'witation t ' . inur limit the rotor diameter. Rotor downvasi and wake effects also are involved because induced velocity is pioprotional to the .•*e'•: square root ef the disk loading. Thus, the higher the disk loading, the higher the induced - or hovering downwash - velocity, which will result in increased ground erosion and greater difficulty for personnel and cargo operations in rotor wal'e areas. Another flcect of disk loading on performance concerns vertical drag, or download. Vertical drag results from the impingemen, of the wake upon the -. fuselage, horizontal tail, and wings (if any). The effect of vertical drag appears as an increment of rotor thrust required over and above the vehicle weight. However, evaluation of vertical drag is not precise. One of the methods described in par. 3-2.1.1.9, AMCP 706-201, employs wske velocity distributions, such as those given in Ref. 2, to obtain dy.
ablcs. The basic analytical procedure for determining
Disk Loading and Induced Power
The relationship between induced power and disk
"5-3
,
-
AMCP 706-202 namic pressure distribt.tions. Drag coefficients are estahlished consistent with the body shapes in the wake, and th-i vertical drag is calculated by a strip analysis. One weakness of this method is the relative inaccuracy of the wake geometry described in Rcf. 2. Improved accuracy of vertical drag calculations is desirable although this mnay require wxtesrsive development of more refined wakc analyses. Model tets can be perfornied with scaled rotor and airframe models. However, Reynolds number effects covninl on data from these tests can be significant, For conventional helicopter shapes (without wings) and values of disk loading, dowrnload is normally about 4-6% (f the vehicle gross weight. Hovering induced power also is affected by blade twist. Tdiij effect is due primarily to altcrptions in spanwise load distribution as a result of twist. I details twist effects for the "ideal" rotor. Twist Ref. seletion for the actual rotor is covered in pat. "-2.1.5. The "swirl", or inplane component of induced velocity is another factor that affects induced power. This inplane component frequently is omitted in the determination of the induced power of the rotor in hover or axial flight. Fig. 5-1. based on work reported in Ref. 3. shows that the swirl velocity effectively rmdu, the magnitudc of the rotational velocity s.cn by the blade element. For lightly loaded rotors, this swirl component can be considered insignificant, but it can be substantial in the more heavily loaded rotors used today. In general, swirl effects should be included in hovering-power-required computations unless disk loading w < 3.5. 5-2.1.2 Blade Loading The thrust produced by a rotor per unit of blade
area is the blade loading BL. This parameter can be defined most sirnnlv in terms of the disk loading w
and the rotor solidity c.
INPLANE COMTO 'lENT OF INDUCEV VELOCITY
VELOCITY (SWIRL) A-VELOCITY SEEN BY OLADE ELEMENT
AXIAL COMPONENT OFVELOC INDUCED ITY (DOWNWASH)
..
Flprt 5-1.
wh-re Rrns R -
V L I EFFECTIVEALL'"ROTAIIONA VELOCITY
--.
..
"
Vector Diagram of Swirl In Hover
mean rotor profile drag coefficient, dii us,
rotor radius, ft
air density, slug/ft! - rotor angular velocity, rad/sec The mean rotor profile drag cocfficient (" is a function of the mean blade lift coIficient CL. in the "ideal" case (Ref. 1) CL = 6Ct° C1 T ( u o.rR~p(flR) (5-3) and C-r a
-
ap(1tR)"
-4)
where
C T
=-
=
thrust coefficicnt, dimension;ss thrust, lb
fiR = rotor tip speed, fps For the more realistic case, where tip losses and other effects are considered, ZrL can be described
where w - disk loading, lb/flt' Srotor solidity, blade area/disk arma, dimensionless More meaningful than this parameter is the niirr, blade lift coefficient eL' This coefficient can be ut-ed
more accurately as 7Cr/c (see par. 3-2, AMCP 706201). Also, a single curve of airfoil seocion lift and drag coefficients cl and cd characteristic of the section is not representative of the actual rotor case, where Reynolds number and compressibility effects are significant. When there are spanwisc variations in blade planform and/or airfoil -'.... *."! actual values of these characteristic coefficients deviate even furthe; from the ideal.
to define the aerodynamic operating point for the rotor blade airfoil sections and, therefore, to det(Tmine the drag coefficient. The profile power required HPo is proportional to the mean drag coefficier,ýif, and can be expressed as
The rclationship between •t and o can be developed from flight tests of rotor configurations similar to the one being designed (i.e., similar in Mach number, twist, and airfoil secticas); or it may be developed from detailed power-required calcu-
BL
HP0
-
-
w o
(5-1)
hp
-,
4400 5-4
lb/ft
(5-2.)
lations that include the spanwise variation of all parameters.
The optimum value of mean blade lift cotfficient
I 9_.CP
06-202
cqLnerally is that value corresponding to (r/UD),.,j,(Ref.4). Further, it is preferable to obtain a blade configuration (planform, twist, and airfoil section(s) ) such that the ratio of section lift 3nd drag coefficients cl/cd is maximum simultaneously all along the blade span. 5-2.1.t Blade Tip Mach NvaSWb Performance and weight considerations generally arm in conflict when efforts are made to optimize rotor tip speed. High tip Mach numbers (greater than 0.65) can be attractive from the points of view of both transmission and blade weight, but they have dctrimental effects upon both power required and noise propagation. If higher tip Mach numbers are employed, tip airfoil selection becomes more critical • hover performance; thin airfoils (thickness less than 10% of chord length) are desirable, and the twist must be selected so as to maintain relatively low tip lift coefficients, 5-2.1.4 Number of Iades I
--
nf
....
unit.s,,C'gr g:ti Inklina
mean
blade lift cocfficient, and blade tip Mach number, rotor solidity has been defined uniquely. With any significant variation from a rectangular planform for the rotor blades, the effective rotor solidity a, should be evaluated using the method of Ref. I. Blade area is defined by the product of rotor solldity and disk area and can be divided among an-, number of blades. Propeller design expeijenc indicates that efficiency increases with increasing numbers of blades. However, recent analytical advances, confirmed by flight and whirl test data, show that tg
-is
t
'r
r..
..
5-2.1.6 Airfoil Sectiom Rotor blade airfoil sections preferred for %lhcir aerodynamic characteristics frequently are incompatible with structural design requirements. and a compromie nmmt be made. In general, for the hovering rotor the inboard airfoil should be of a lowdrag type (at least with extensive lower.surface laminar flow). Outboard of 70% radius, compressibility effects i..ust t a considered, and the lift-to-drag ratio L/D for the airfoil section slould occur at the local Mach number and angle of attack. These conditions suggest a spenwise variation in airfoil contour. If a constant airfoil is employed, its scleclion should be weighted toward complying with the angle of attack and Mach nun.ý,er conditions at or near the blad- tip (outboard of 80% radius). 5-2.1.7 Hovering Thrust Capability The capability of a hovering rotor to produce thrust can be expressed by a simple relationship. However, the agreement between the calculated and measured values of thrust produced for a given amount of power applied to rotors of practical cons are ,prove•ne, figuration is not good. Several " available and arc reviewed in par. 3-2.1.1, AMCP 706-201. The method most appropriate for calculatinS the capability of a new rotor possibly is dependent upon the similarity to rotors for which analytical and experimental results are available. The limitations of tOie available methods for prediction of the performance of hovering rotors also is discussed in Ref. 5. 5-2.1.8 GuIdeflin o 'rthv.
.u-*v.rn
Apparently, intcrblad¢ interference can reduce the hovering efficiency of muitibladed rotors (Ref. 5). The selection of the number of b!ades, therefore, is dependent morm upon considerations of overall rotor system weight than upon aerodynamic efficiency (see
eters requires systematic parAmethc varip.tion involving all of the major variables given previously. This analysis is discussed in detail in par. 3-4 1, AMCP 706-201. Generalized results arce given in the paragraphs that follow.
par. 3-4.1, AMCP 706-201). S-2.1.5 Twig Selection of blade twist for the "'ideal" rotor is covered in Ref. I in current helicopters, twist
In current helicopter designs, disk loading generally does not exceed 10 lb/ft' . Light helicopters (less than 5000 lb gross weight), tend to have disk ioadings of 3-5 lb/ft2 . The medi in,-size helicopter, 500015,000 lb tends to be in the 6-5 lb/fl' class, and for
generally is linear in order to simplify manufacturing. If stretch-formed spar3 are used, nonlinear twist isobtained quite easily. In any event, twist selection is a function of disk loading and blade tip Mach number. The higher the disk loading, the greater the optimum twist; and the higher the tip Mach number, the greater the required twist. Twist optimization is achieved by systematic variati.ns using detailed analytical methods.
larger helicopters disk loading is of the order of 10 lb/ft' . Si'e and weight effects bias the disk Ioadings higher as gross weight increases. The curremt cinphasis on high-altitude, high-temperature design conditions results in values of mean b!ude lift coefficient ?'L values of the order of 0.44 to 0.54 for sea level standard day conditions at primary mission gross weight. Current helicopters have hove:ing blade tip Mach 5-5
numbers rangting ftom 0.50 to 0.75. Weight and struc-
tural conciderationh suggcst higher minimum valuts, and noise considerations suggtst lower maximum values - resuliOi in a conpr'nise design fange Iyetwecn Mach 0.6 and 0.7. 5-2.2 HIGH-SPEED LEVEL FLIGHT To maximize bigh-tpeed level flight performance, the same patr metta are considered as in optimizing hover performance. in addition, the ratio of flight hspeed to rotational tip speed, or advance ratio o, is intraduced. In high-seed . design, the basic compoatase between advancing blade tip Mach numberoad advbnce ratio. The advancing tip Mach number Man , can be defined ts
cannot be divorced frori the hover and maneuver requirements. However, it is discussed as x separate probklm here, where for a given amount of power available, airpeed is to be maximized, Initially. a source such an Ref. 6 crn be used to determine an initial set of values for twist, solidity, and tip speed. This source requireiy that values for gross weight and vehicle parasite d;,ag area first be assumed. Ref. 6 also assumes a particular airfoil section and a linear twist distribution. From this starting point, modificationt of blade tip airfoil section, planform shape, and twist can be made in order to achieve speed increases up to the limits of the power available. To increase the advancing blade tip Mach number at which drag divergence becomes critical, airfoil thickness can be reduced. For symmetrical airfoil&,
reduction of thickness to vaiues of les than 12% norV + RR ad
I#,,
a
(5-5)
where W speed of sound, fps a - true airspeed, fps V At a given forward speed, decreasing tip apeed d __ecras•s the amount of bUia, that i providing us•e uI lift and propulsi've force, because more and more of th, 411 is in reversed flow. This effect is accompanie&,, nocessarily, by increased lift coefficients over which eventually can of the disk, part amounts the of stall. lead"working" to signi•ficant
The alternative approach is to increase rotor tip speed. This leads to increasingly higher advancing blade Mach numbers. Eventually, drag divergence is attained over a significan. portion of the advancing blade, with increased power requirements a& a rsult. Increasing blade area with a given value of rotor
tip speed will lower the mean blade lift coefficient
and, therefore, allow operation at higher advance ratios. Increasing twist tends to alleviate the retreating blade stall problem up to a point, but also can result in negative lift on the advancing blade tip. The latter is disadvantegeous because higher lift coefficients must be achieved over the positive-lift portions of the disk in order to compensate for the negative lift on the advancing blade tip. Also, with large amounts of blade twist, drag divergence - with an accompanying increast in power required - may occur due to high negative angles of attack on the advancing blade It is necessary to determine the cormbination of tip speed. solidity. and twist that results in the minimum power required for a given speed, or the maximum sp-ed for a given amount of power available, Normally•, the high-speed performance problem
5-6
mally results in a reduction of maximum lift coefficient. This is detrimental for thet lifting capability of the retreating blade. This effect can be altered by introducing camber into the airfoil scction of reduced thickness in order to maintain an acceptable value for C,,,. while also attaining an increased drag diverof _nt, i-_,Ip_.m, ,_,n 1_r... -. gpnp P _=h .--,-m camber will result in undesirable blade pitching moments at high level-flight Mach numbers. Sweep of the blauc tip can be employed to decrease number, thus allowing higher the effective Mach values of rctual advancing tip Mach number (V +I
OR)/a before the drag rise due to compiassibility becomes unacceptably high. However, care must be exercised to avoid the loss of effect-ve area and, therefore, of retreating-blade lift capability. Nonlinear twist distributions may assist in optirizing speed for a given amount of power available. i6i det£iled Ioto a iTA CzIVts IIIUhM Ve invitIed analysis by consideration of radial and azimuthal variations of angle-of-attack and Mach number. No r, Its can be offered; trkl-and-error is the only approach currently available. 5-2.3 HIGH-SPEED MANEUVERING FLIGHT Achievement of the desired maneuver capability at a given airspeed also may affect tat selection of final values for the basic rotor design parameters. Because of increasing amounts of retreating blade stall, tht higher the forward speed (for a given tip speed) the more difficult it is to achieve high maneuvering load factors. To begin with, a static analysis is not satisfactory for determination of maneuvering flight capability. As discussed in Ref. 7, rotor pitch and roll rates are involved in both symmetrical and turning maneuvcrs, and can affmt load factor capability signifi-
AMCP 70t-2022
• J-.
cantly. These maneuvei rates aler the cnglo-of-attack distribution obtained during steady-state flight at a givtn speed and rotor thrust level. In general, to a,;hieve high maneuver capability, blade loading in trimmed steady-state flight, i e., normal load factor nZ - 1.0, must be low. Load factor, or maneuver, capability can be related to (CT/ra), ,/(CT/u)Iwz - 1.0. Thus, for a given rotor design with known (CT/c),,, the lower the trin thrust cotfficient (or blade loading), the greater the load factor capability. The other -,:rn under decgn control is (C'/),. The major variables for maneuver capsbility are advance ratio, airfoil section, and twist "or a given solidity ratio. decreases As advance ratio increases, (CT/O)W(Rcf. 8). Therefore, for a given flight speed, an increase in rotoi tip speed increases (Cr/c),,. However, as for level flight, a maximum value of Pdvancing blade tip Mach number must not be exceeded, Advancing blade shock stall can be encountered if the Mach number is too high. The magnitude of (CT/i)6 for a given advance
tational kinetic energy KER is defined as A (S
KER
I
I
) ,
-
(0-o)
fi-lb
where mass momcent of inertia of the rotor. slug-ft2 The symbols fl,,,i ad fl,,d represent the rotor angular velocities at the beginning and end of the flare maneuver, respectively. However, the determination of an acceptable value for (fi,1 for a new rotor is largely judgmental, with little more than the designer's experience available to assure that the rotor remains controllable throughout the flare. Computation of helicopter autorotativc performance is discussed in further detail in par. 3-5.1. AMCP 706-201. In par. 3-5.3, AMCP 706-201 an autorotativc index AI is developed. Acceptable values of this index, and henctc of the rotor inertia, also are discussed. 'R
-
ratio is a strong function of the maximum stetion lift coefficient near the blade tip. This is not ne.essarily a dircta funct-so of the sction, or two-dir,,-enioal, maximum section lift coefficient cl., because of complicating factors such as spanwise flow and oscillating airfoil effects. However, it is a good general
5.3.1 GENERAL Rotor systems can be described as articulated, gimbaled (or teetering), hingeless (sometimes referred to as "rigid"), and flex-hinge. The blades of an articulated rotor system are at-
of the tip section will
tached to the hub with mechanical hinges, allowing
improve the rotor maneuvering theust capability, Because retreating blade stall generally occurs first, the magnitudis of cl.1 . at the retreating blade tip
the blade froedon, to flap up and down, and swing back and forth (lead and lag) in the disk plane. The blades of the hingeless rotor are attached to the hub
rule that an increase iii c,..
Mach number also is quito important. New airfoil design developments (Ref. 4) allow a tailoring of the at a sction profile to obtain the peak value of dosirod Mach numbr. Blade twist also affects maximum thrust capability
without mechanical hinges for flapping or lead-lanotion. The flex-hinge, or strap-hinge, rotor enploys a flexible structural attachment of the blade to the
by controlling the lift distribution at the retreating
ness of the articulated or gimbaled system.
h
-,,h!uh,,,•,,,•u
.
Generally, the type of rotor system will have been selected during preliminary design. In par. 3-3.3. AMCP 706-201, each of the types of rotor system is
mization of twist for the maneuver case usually is mdetrimental to level-flight performance, of a omprot-
dcribed, together with the methods by which each is controlled and in turn is used to provide control of
miae o.een ig required. Computation of helicopter
the helicopter. The discussion includes a simplified
detail in par. -. , AMCP 706-201.
articulated) rotor, while the dynamics of rotor
5-L4 INERTIA Rotoratertics
summary of the flapping motions of a flapping (fully
systems are described in detail in Chapter 5, AMCP
•", erain autorotative
106-201. The descriptions of the several types of rotors in par, 3-3.3, AMCP 7dti 201, include dis-
landing characteristics. Rotor angula velocity and
cussions of the advantages and disadvantages of
inertia uniquely define the rotational kinetic energy of the rotor that "an be used in the development of a decelerating force to arreqt descent velocity in a zero-
each, together with a review of the heiicopter siLes for which each may be most appropriate. The discussion of rotor systems kinematics and
power or partial-power landinli. The amount of ro-
controi which follows supplements the introductory
Ro
t
IS,
.,. IL,,,
high stiffness of the hingeless rotor and the low stiff-
blade tip. Optimum twist is determined only by detailed analyses of the maneuvers including major effects such as pitch and roll rates. However, opti-
is discussedin more erformanM C 3 ancuvering flight
-
hub
0.
5-7
'
AAMG 71*202-
TYPE VECTOR TILT
MOMENT SOURCE
LIFT VECTOR
FLAP
1. THRUST VECTOR TILT 2. HUB MOMENTS DUE TO SHEAR FORCE AT HINGE
(A) ARTICULATED ROTOR
LIFT VECTOR TILT"
L.-
I
TU
V
I
\ROTOR
V1
SRTILT
\ANGLE
(B) GIMBALED OR TEETERING ROTOR
LIFT VECTOR_ VECTOR TILT /
i
1. SMALL THRUST VECrOR TILT
IHINGE
2. HUB MOMENT DUE TO SHEAR FORCE AT EQUIVALENT HINGE 3. HUB MOMENT DUE TO BLADE
STRUCTURAL STIFFNESS
(C) HINGELESS OR FLEX-HINGE ROTOFR. Figur 5-2.
Contlrol Mommil for Bai
Rotor Ty.pes
AMCP 706 202 ,decription in par. 3-3.3. AMCP 706-201, and the 5, presntation of Chapter theoretically oriatted
LAG HINGL
AMCP 706-201. PITCH
$-3.2 HELICOPMTR CONTROL Inflight control of the helicopter, using the rotor types cited. is provided by: 1. Momenta acting upon tho rotor hub 2. Tilting the resultant rotor lift vector 3. A combination of these. I he control moment source for each type is illustrated in Fig. S2. Fom the gimbaled rotor, a given rotor tilt produces a corresponding tilt of the lift vcctr, which, in turn, produces a control moment about the helicoptcr CG. An additional control mon~elit exists in an articulated rotor as a result of the hub shear force acting at the flap hinge to produce a moment at the hub. In the case of the hingelcss and flexhinge rotors, the structural spring at the equivalent hinge provides an additional component of control moment et the hub. The conventional metbod of achieving rotor control is through collective and cyclic pitch changes at the blade roots. These changes arc accomplished
freedom of the blade is necessary in this particular design so that the steady chordwise bending moment at the blade root is reduced. In the cquilibrium lag
throiugh cntril Ignkagc between the rotating blades
pe.t~inn of the blade, the chordwise moment due to
and a swashplate (a structural lecmcnt thai constilutts a fixed plane that defines the blade pitch as a function of azimuth). Individual blades arc mounted on spindles that provide feathering freedom for control. Collective pitch of the blades is introduced by a scissor niechanism or by raising or lowering the swashplate; cyclic pitch, required to produce a tilt of the rotor disk plane, is accomplished by tilting the swashplate. Blade pitch changes also arc made in some rotor systems by connecting the swashplatc to a servo tab or. ... or kc. ... . .......... .. . ..... oa servo rotor or gyro bar that in turn acts as a swashplate (or the main rotor,
the drag loads on the blade is balitcd at the lag hinge by an opposite moment due to the centrifugla force and the lag displacement of the blade. in-
5-3.3 ARTICULATED ROTOR The kinematics of an articulated rotor with an outboard lag hinge are illustrated in Fig. 5-3. Veitical motion of the pitch link in response to swashplate tilt as the blade travels awound the azimuth produce, pitching rotation at the pitch bearings corresponding to the cyclic pitch of th-. rotor. The position of the pitch bearings with respect to the blade lag freedom varies with the rotor system design. In the example (Fig. 5-3), the pitch bearing of the rotor system is inboard of the lag axis, whereas that of the CH-46 rotor is outboard of the lag hinge. As shown in Fig. 5-3, the lag hinge allows the blade to move, leading and lagging, in the disk plane. Lag
LAG DAMPER N
-
HINGE
,1l, HUB ROTATION |;igrc 5-3.
Artlculaled Rotor Schematic
dividual blade lag dampers are required to provide energy dissipation adequate to conrol the mechanical instability associated with the coupled rotor/airframe system as dcacribcu in par. 5-4,3 (also see Chupter 5, AMCP 706-201). The rigid-body lag natural frequency of articulated rotors usually is betwoen 0.20 and 0.40 times the rotor speed. Blade flapping freedom in the articulated rotor is ro-d-d ky. a h r rnia i me, which is located close to the rotor cntcrline in order to minimize the flap bending of the rotor hub (Fig. 5-3). The steady ioo"mentabout the flap hinge from thr centrifugal fore acting through the moment arni of the blade - vcrtically displaced by the blade coning above the dirk plane - is balanced by a moment of the same magnitude. b':t in the opposite direction, due to the steady lift on 4he blade. The natural flapping frequency of an articulated design isnear resonance with the rotor shaft speed. However, aerodynamic damping in the rigid flap mode approaches 50% of critical damping. with the ruult that the near-resonat conditioa provides an acceptable design. Coupl;ng between flap and pitch motions is &n important dcsigrn consideration for a rotor control systkm. Generally, the rotor should be designed so that, as the blade flaps upward, the mechanical pitch 5-9
J 1
,'
"AMCP 706-202 angle of the blade remains the same or decreases. The kinematic coupling that varies the feathering, or pitch. angle of the blade with flapping is defined as 63, and the standard notation is that an increase of pitch with an increase of flapping angle is positive. Flappitch coupling can be introduced mn!chanically by a skewed flap hinge, or by radial location of the attachment of the pitch link to the pitch arm inboard or outboard of the flap hinge. Negative 63 generally is required to improve stability of the rotor (see Chapter
N AM
5, AMCP 706-201).
Pitch bearings outboard of the lag hinge produce . kinematic oupling that changes the blade mechanical pitch aihgle with blade lag motion. This conflijaration has the potential for unstable pitch-lag blade motion. Fig. 5-4 illustrates the general arrangement for a rotor with coincident flap and lag hinges. This rotor has a compact arrangement of &ap and lag hingcs exactly like P universal joint. Flap hinges located further outboaad provide greater control power, but also increase the flap bekiding moment at the hub. The location of the lag hinge closer inboard results in a lower lag natural frequency, with increased damp-
PITCH
FLAP HINGE HUB ROTATION Figure 5-4.
Coincident Flap and Lag Hinge Rotor
PITCH ARM
aing heing required to prevent ground resonance.
5-3.4
GIMBAi ED (TEETERING) ROTOR
Fig. 5-5 provides a schematic of a.gimbaled rotor
system, only two blades of a four-bladed rotor are shown, although any number cLf blades may be used. Each blade is mounted on a spindle attached to a yoke that interconnects the blades. The yoke, which "definesthe rotor disk plune, is ginbal-mountcd to the helicopter mast (the top of the rotating shaft). In a gimbaled rotor, no cyclic pitch motion of the blades occurs relative to the spindles for any steae; hovering condition, regardiss of CC, location or flapping rclative to the mast. The phase relationship of the orie-per-rev excitation of the primary inplane bending mode is such that the blade root moments are reacted internally in the yoke, leaving the rotor hub undisturbed. The yoke structure must be stiff enough that the natural frequency of the blade cantilever mode is sufficiently greater than the rotor speed to avoid excessive amplification of one-per-rev loads. For two-bladeci, or teetering, rotors (Fig. 5-6), the gimbal mounting of the blades may be replaced by a single tectcring hinge that allows only seesaw or flapping motion of the blades. Cyclic and collective blade pitch oc :,rs about the yoke spind!es. For rotor tilt relative to the' shaft, the blades are forced by the trunnion out of their ideal position in the cone of the rotor twice each revolution. This results in a bending 5-10
0
"
YOKE YOK)
Flgure &-5. Gimbaled Rotor Schematic moment on the mast in the direction of rotor tilt that varies at a frequency of two-per-rev. A sketch of a teetering rotor system is shown in Fig. 5-7. This rotor is connected to the shaft by a hinge, the axis of which passes approximately through the CG of the rotor in order to minimize vibratory hub and control loads. The stabilizer bar prorides stability by increasing the lag time between shaft tilt and rotor tip path plane tilt. T he stabilizer bar is connected to the blades through mixing levers bctwecn the sides of the bar. The inner ends of the mixing levers are pinned to the bar, the outer ends are
I-AMCP
W0
VF YOKE SPINflI E TEETERING HINGE
FLAPPiNG bF5gure 54G.
Teetrlng Rotor SEd'.cma2tic
MIXING LEVER
following rate and imnproves the maneuverability, but also degrades the stability.
~
:
HINGELE~SS ROTOR blade of a bingclcss ~otor is illustrated scheniaticalty in P~ig. 5-8. In this type of system, no sme!~ ~ STABLIZE BAR chanical mcawis are provided to allow chordwisc or1 flapwise displacement of the blades. The blades are UAMPLN rf dhfror the rotorfhub. which is attached rigdlyto he otaingshat.Collective and cyclic SHAF1pitch inputs for variation of thrust and control moPITCH ARM ment are made through the pitch links in response to pilot input to the swashplate. The pitch angle is 7 f -6)changed by rotation of the blade about the feathering axis just as is an articulated rotor. Following a cyclic pitch input, the hingoless rotor responds as shown in Fig. 5-2(C), providing a control moment about the helicopter CG as a result of both tilting of the recultant lift vector and a moment acting at the hub. natural frequency of the first flapwise bending 5-. RtorThe Figue Teterng Ilgue Teterng 5-. Rtormode fixes the offset of the equivalent flap hiinge. The dynamic characteristics, control power, and pitch and roll damping for a hingcless rotor are identical to Sconnected lo the swashplate, and the middle is cont'osc of ani articulated rotor whose mechanical flap mcted to the pitch arms. The damper regulates the nlgc is located at t~he equivalent hinge point. A rate at wHich the stabilizer ha.- follows the tilt of the t.Lnrgeless rotor with a fundamental flap frequency of rotor shaft. An increase in damping quickens the between 1.10 and 1 15 times rotor speed would have UNIVESAL OINTOne
_
)
-
5-3.5
AMCP 70F-202
S~&
'
CYCLIC PITCH
./-
S~PRECONE/
ROTOR HUBA
Figure 5-8. Hingeless Rotor Schematic characteristics similar to those of an articulated rotor with a flap hinge located at 20% of the blade radius. Precone of the blades, typical of the gimbaled and hingeless designs, permits cancellation of the steady lift moment by the moment due to the centrifugal force of the rotating blade acting through the vertical disolacemcnt of the blade above the disk niane. Kinematic coupling in a hingeless rotor is influenced by the location of the feathering axis with respect to the precone angle and to the equivalent flapping hinge location.
S5-'
5.3.5.1 XH-51 Rotor System A schematic of the Xh-51 hingelcss rotor, which has cantileverd blades with only a feathering degree of freedom, is shown in Fig. 5-9. A mechanical stabilizing gyro is connected by one set of links to the blade pitch arma and to the rotating swashplate by another set.. Blade cyclik pitch is controlled by the control gyro, which, in turn,, is controlled by the swr.shplate input. The blades of this rotor have high chordwise stiffness, but arn provided with flap flexibility by a flat spring section inboard of the feathering axis. The blades are swept ferward about I deg ahead of the featherin, axis to locate tme CG of the 2
blade ahead of the feathering axis. Thus, the inertia forces acting theough the CG produce moments about the feathering axis, producing, in turn, feedback forces at the gyro. Analysis has shown that this displacement of the CG forward of the feathering axis permits the gyro effectively and simply to provide stabilizina pitch innuts to the rotor. In this system, upward flapping of the blade puts an up force on the pitch link, causing the control gyro te precess. ThL. tilting of the gyro then puts cyclic pitch back into the blade at the proper phase to minimize blade flapping motion. The control gyro also provides the blade pitch angle changes necessary for stability of the helicopter. The AH-56A rotor also is of this type, but us= a door-hinge arrangement of thu pitch change (feathering) bearings in order to obtain the desired high chordwise stiffness of the hub without excessive drag. 5-3.5.2 OH4A Rotor The OH-6A rotor syst:m, shown in Fig. 5-10, is another type of hingelcas rotor. Multiple straps transmit the centrifugal force from one blade across the hub to the opposite blade, and are flexible enough to allow both flapping and feathering of the blade.
*AMCP
70t=
CONTROL GYRO
NOATON-\ OTTNG. WSHLT
RESULTIOGATLIGhT DIRECTIO
Vlgre 5-9. XH-31 Rotor System Curved thous installed at the point., at which the h.4. Vu Wt"No "ov S&Arape ar W i. inS of the straps at any one point. In this otherwisic hingekas rotor, lag hinges are located at the outer ends of the tenaion-flati-torsion straps. Excessive static droop of the blades is pre~vented by stiff cuffs that are attached to the bilades and cover the straps. Co.itact between the inboard ends or the cuffs and the hub lim~ts the downward and upvarJ flapping excursions of the blades. ROTOR SYSTEM KINEMATIC COUPLING Adverse kinematic coupling can result in various types of instability in a particular rotor system. This paragraph reviews tlac subject independently of the rotor type, but considcnt the blAdv and its retcnt.on system. The niechanism of rotor instability resultivg \ from blade kinematic: is examnined we~akratcly undet 5-3A6
)
thec cate~gorie* of pitch-lag, pitch-flap, and flAp-lag. n
Iut~ in " ivtientn fi-
nf th
ionatahilitiem is
outlined in Charter 5, AMCP 706-201. $ .3.6.1 Pitci-Iag Iftability Rotor blades with substan~tial chordwisc displace metiv have &potential "pitch-lag" instability. The critical degrees of fr'-cdom involved are fiap and lag. Howtver, the critical design parameter is a kinematic courpling that causes a blade pitch angle change in rzponse to lag motion, or -hordwise displacement. The mechanism of this instability is depicted in Fig. 5-11. As the b'ado lags (A). and if the pitch-lag coupling causes the blau1c pitch to decrease (b). there is a loss of lift. Downward flap of the blade occurs due to lift loss (C), and produces a Coriolis force in the lag direction (D), causing additional blade lag. Further discussion of this phenomecnon may be found in Wtf. 9 and 10. 5-13
•
.
-. ;< U. .. i.• . . .. .;'
I
~k" -
...-.-_-
____.-.
"LAGGING
"-.
~~~~LIMITS
(FLAPPING
STRAP RADIUS S.VOE BFNDING BEND
•
FEATHERING
".••• ,(
Figure 5-10. OH-6A Hingelem Rotor System
"5-3.6.2 Pitch-flap lstabllity Rotor blades are subject to the anme sort of dynamic instabilities as are fixed wings. For example, they are susceptible to the classical bending-orsional flutter discussed in Ref. 11. For hover or vertical flight, the major difference between the rotating and the fixed wing is the velocity variation spanwise along the bWede due to rotation. The principal paramtPre
inflpnpring thie mnjie
in hnth euetpms arm. th@
chordwise distapce between the CG of the airfoil sec. tion and the aerodynamic center, and the torsional stiffnes, In addition to to.sional deflections, either flapwisc or chordwise displacements of the blade deflections also may interact in such a way that a pitch-flap instability can occur. The case of a blade with flapping deflections above the feathering axis (flapping hinge outboard of the pitch bearing) is illustrEted in Fig. 512 (A). As the blade flaps with respect to its steadystate position, the resulting Coriolis force produce4 a pitching moment about the feathering axis. If there is flexibility in the pitch control system, this pitching moment caue the pitch of the blade to ;hang;. Therefore, stiffness of the control system also is a significant factor in pitch-flap stability. As shown in Fig. 5.12 (B), the same blade section 5-14
initially is at distance r sinpo from the feathering axis. As the blade rotates noseup about the feathering axis, the blade section and the lift force acting upon it are displaced backward Or sini 0 , producing a noscdown pitchin$ moment. This moment, together with those caused by variations in the inplane aerodynamic force and the centuifugal force acting through the moment arm r sin~,results in coupling between the nitch and flan dearees of freedom and; conseuently; affects the stability characteristics. The net pitching moment about the feathering axis changes the blade pitch angle, hence the angle of attack, by an amount that is inversely proportional to the control system stiffness. Completing the cý :4: for pitch-flap motion, which may be unstable, thv blade section lift varies as a result of the tngle-of-attack change. The lift variation causes blade flapping, which, in turn, produces additional Coriolis forces. Steady inplane bending deflection or bladt sweep also can introduce pitch moments as a result of lift variations. 5-3.6.3 Flaplag Instability Ref. 12 describes flap-lag instabilitizs as a result of finite blade deflections. The conclusion from this werk is that lifting rotors that have no laS hinges iniy, under certain conditions, be subject to limit-
,,
,
• .• ,
J-
7'
(A) BL ADE LAGS
(B) PITCH ANGLE DECREASES
(C) BLADE FLAPS DOWN DUE TO LIFT LOSS
(D) CORIOLIS FORCE CREATED INLAG DIRECTION PRODUCES MORE LAG Figure 5-l1.
*tional
Mheanism of Pitch-lag InslabIllty
cycle instability in both vertical and forward flight. The basic mechanism of this type of instability in. volves the steady-s' ite blade con'ing angle. For a blade with positive t 4kl of attack, the local wind yelocity is increased by the blade lead velocity, rebulting in an additional lift, as shown in Fig. 5-13 (A). The resulting incremental flap-up aerodynamic moment is counteracted by an equal flap-down centrifugal forme moment. The incremental centrifugal force also Jresults from the lead velocity or incremental rotavelocity (Fig. 3-13 (3)). If the steady coning angle is obtained by the balance of thrust and centri-
fugal moment, thewe two opposing incremental effects arc equal, and no coupling exists between flap and lag. However, if the coning angle is reduced because of elastic slapping restraint, the vertical comnponent of the centrifugal force vector is reduced and the incremental aerodynamic flapping moment exceeds the centrifugal restoring moment. In this cane, the flap-up moment produced by a lead velocity of the blade can result in unstable blade motion. In forward flight, an additional destabilizing turn occurs. This is an aerodynamic flap moment proportional to the product of mean lift, lead velocity, ad-
.
HIORIZONTAL
A0OITIONAL
FFORCE MOETMU4
LIFT SECTION LIFT
0SUTJJ LOCAL. WI40
rC"'LEAD,
VELOCIT INFLOW
NANG CENTRIFU
\
(A) ~ ~
1
~nP
OF ATL
T,
I)
ENTREFUGALOFORC DU B ~ COILNCRENMRO ~LAEFLPDENTOLAL ~~ LEAD VELOCITY
neve Sibainntespa fsiunie) d/til s#OFs and31a stffes oFORCExsas-
(B) ~~FOROM ~ ~ LI FT~OFSTBL.Hwvr ~ ~ ~
t
F~ure5-1.
minif
~ PICEMCHA ~ B MOMENTUGA tercnieainegsAtcDE-
Pld~-lapCou~l~ug o Roors
vanoess the sinele ratio ofpe th azimuth andfneane.A u OEN blade moFTi OF anEoccurt
haroni PITabiiH moderaeadva2. pitihfo~. p Cop
EETLDE
ortype ofinling0.4. Thi
flection aotd grossg viraoryidresponse ar becmias faily wpellr debaned ond calulaion profedues arema aevilablequnisdtiso.baigsrspa
3-*
Hoetr y otherCU considerations eOf. s~tatc i
andbrat
Ladsmambeomn ibaor y
stability can be suppressed by an adequate amount of inplane damping. Ref. 13 suggests that, if the blades are sufficiently flexible in torsion, pitch-lag coupling can be used as an additional means of suppressinS lag motion instability.
The procedure for development of preliminary rotor design into a detail design requires continuous iterations by aerodynamnics, dynamics, stres, and fatigue specialists. The integration of computed vibratory rotor loads into the detail design is dcscribed best by a generalized mothod toward which the in-
5-4 ROTOR SYSTEM DYNAMICS
dustry is moving.
5-4.1
OSCILLATORY LOADING OF ROTOR BLADFS The o lsciltory loading of rotor blades is important for both vibration and fatigue. Rotors usually are deSgned by extraapolation from previously successful designs. This can be. expensive because many rotors have been designed, bui~t, and flight Stetd without significantly exteiding the useful envelope of existing rotors in relation to loads and vi5-16
Ideally, the following program, which hinges on the existence of a complete seroelastic flight-vechicle analysis, could be used to link the preliminary and detail designs: 1. A nmaneuver spectrum should be defined, based upon previous helicopter experience and the mission requirements of tie now helicopter. 2. The maneuter spectrum should be "flown"' analytically with the prelimninary design helicopter, and loads and vibrations should be computed at the
• 1•
) -
locations where flight test inhtrumcntation will be placed. Thesc calculations should cover the complete ground-air-ground cycle, including landing, taxiing, and towing. In such a sequecnce, all of the static, transient, and oscillatory stresses and vibrations should be compasted. 3. The stresa and vibration calculations should be reviewed to determine where design optimization should start. If the performance, stresses, vibrations, handling quaaities, and other rotor characteristics meet but do not exceed the specification requiremeats, then the preliminary design may represent an optimum configuration. if the helicopter greatly exceeds requirements in one or more of these arcas, a weight penalty generally results. Design changes should be made and at least a portion of the analysis based upon the maneuver spectrm should be recomputed. If tht design fails to meet the specifications in certain arew ,, the design must be improve I 'nd a portion of the maneuver spectrum analysis should be rccomputed to demonstrate compliance, At the end of the iteration, thi detail design should start with a sat of loads for static and dynamic stress calculation, and a reasonable assessment of vibra, tion The detail design of the rotor system should be initiated by making dimensional drawings of all parts. The designer should follow the preliminary design as closely as possible. However, compromises often may be required, some of which can aiter the dynamic chepracteristics significantly. As the design progreases, section properties, weights, inertias, and stresses should be computed for the detai! parts. As soon as the first design iteration is completed, the rotor natural frequencies, loads, and stresses be recomnuted. These stresses should be used should as the .basis for a fatigue analysis, which should be guided "additionally by experien= from the test histories of it'milar parts ana idealized material samples. Depending upon the outcome of this first design iteration, the design should be accepted or another iteration started. Significant advances have been made with helicopter flight simulations. For example, Ref. 14 describes a recently developed analytical tool for computer "flight testing" of VTOL designs. Steadystawe flight, maneuvers, and gust response effects are included so that required cornfiguration changes can be made readily during the preliminary and detail design stages. This objective is achieved by detailed reprewentation of the aircraft, including rotors, wings, auxiliary propulsion, and control systems. Complete blade element analysis of the main rotor(s), tail rotor, and propeller(s), as applicable, are performed
AMCP 7W6202
through a maneuver, and are based upon the Instantaneous aerodynamic and dynamic environments. Time histories of rotor blade loads and bendins moments are calculated. The basic equations and programming procedures are presented and discussed in Ref. 14. The representation of airframe rnd rotor parameters, the types of maneuver inputs, and the available output formats also are discussed in detail. Typical case studies are given. This analysis is capable of evolving into a generalized procedure with the additiou of details such as elastic pylon and fuselsge, and a fully aeroelastic rotor. Complete documentation of this particular method can be found in Ref. 15. Comparable met':ods have been developed by other contractors. 5-4.1.2 Oscillatory Load Design Comiderations Oscillatory loads are a major factor in rotor design; but the calculation of oscillatory loads is not yet sufficiently accurate for life prediction and design assurance. Therefore, rotor design is guided by calculatcd natural frequencies, static loads, and factored oscillatory loads. The final demonstration of design adequacy comes from flight and fatigue testing. 5-4.1.2.1 Rotor Oscillatory Load Calculation Most current procedures for computing rotor natural frequencies and loads are based on Myklestad's development of the dynamics of a rotawing beam (Ref. 16) and on simplified, two-dimensionl aerodynamics (Ref. I). Typically, such analyses can 15 used to compute natural frequencies and airloads separately; ther the two analyses are combined to compute the forced steady-state response. Many versions of this procedure have been developed. A rit!c ecito foeaattowi.hhos been used for designing two-bladed rotors for nearly a decade, is given in Ref. 17. 5-4,1,2.2 Drawing Board Phase As noted previously, the drawing board phase of the rotor detail design is an iterative procedure and typically involves several groups of engineering specialists. The procedure is best explained by briefly describing the functions of several elements: I. T'he areodynamics group sizes the rotor, de',clops the blade contours, and helps determine the static load spectrum during the preliminary design phase. Often this work carries over. with little change, into the first step of detail design. 2. The rotor design group starts the design iteration by laying out the preliminary rotor design and developing dimensional drawings for the detail parts. The designers must be cognizant or the stresses and 5-17
-
*
•,.
.1.
vibrations resulting from rotor oscillatory loading, Other considerations requiring attention indude dynamic stability, wught control, bearing applications, mechanical function, value analysis, materials, bonding, and manufacturing processes, and the designer must rely upon specialists in many of these fields. 3. The rotor stress group begins preliminary design with the development of section properties, These properties, in turn, are used for determination of the nominal stresses that result from the highest combinations of cenuifugal force (CF) and maneuver thrust )onditions. The bending .Ioments for these conditions are computed in the steady-load portion of the rotor-load analysis (par. 5-4.1.2.1). Fatigue may be considered by an empirical method that provides a simple and reliable account or load spectrum shape and severity in relation to componerat fatigto strength in the initial stages of rotor component design (Ref. 18). This method uses the flight :oads calculated for maximum level flight speed, the condition that usually produces the highest continuous (nontrans~eut) alternating loading. A factor is applied to the calculated loads to obtain design loads that will oroduce a satisfactJev fatiaue-life structure. This factor is a function of the material SN curve, the maneuver spectrum severity, the loading frequency, and the fatigue life required by the helicopter system specification. The design curves for this factor are based on an analysis of several flight load surveys. Natural frequencim also are calculated for the rotor as defined at this stage of the design. Changes in section properties and/or concentrated weights are used to produce a frequency distribution that avoids principal resonances. I~~ltn-;.• eh.• *•• A-.t;I A -. __L-.SIR. U%,A
Z.--66,,
P,,OR%,
r rotV o
natural frequency placement insures the lowest possibk oscillatory rotor loads; and rotor and fuselage frequen.y placement, isolation, and superposition in"sur the lowest possible fuselage vibration. The rotor characteristics must be such that low-frequency vibrations of the fuselage are avoided. High-frequency vibrations generally are not so critical and may be corrected during the flight test phase. The aim of this continuous iteration is a design optimized, or balanced, with regrd to performance, function, strength, life, weight, and vibration. When an acceptable balance has been achieved, the design is considered to be adequate, the drawings are compked, and rotor components are manufactured. However, further changes usually result from the
flight and fatigue tests. 54.1.2.3 FM TeMt ,•
The design calculations made during the drawing . -1S
board phase incure that a rotor mects the static strength criteria and has an infinite life for the low. cycl, high-stress variations associated with the ground-air-ground cycle. However, because of the superimposed high-cycle, low-stress os.ilations, mcny rotor and control system parts will have finite fatigue lives. Design changes often are required to insure that the adp~luate component fatigue lives of all components are adequate. Flight and fatigue testing is required to determine these lives. Prior to the flight test phase, the strain gage instrumentation, the flight load survey tests, and the data reduction format shall be specified. The measured loads are used, along with the approved maneuver and frequency-of-ocurrence spectra, to determine the fatigue lives of all fatiguc-critical components. The required flight load survey tests are outlined in Chapter 8, AMCP 706-203. The calculations made during the drawing board phase are effective in guiding the design so as to avoid excessive loads and vibrations at the lower frequencies; however, additional tailoring of the rotor and fuselage usually is required during flight test to minimize high-freouencv loads and vibrations. Measures taken include optimizing the amount ard location of the concentrated blade weights; changing blade stiffness, such as through use of trailing edge stiffeners; optimizing the pylon suspension pa- i. meters, and detuning the fuselage by varying stiffness and the location of certain concentrated weights, e.g., the battery. 5-4.1.2.4 Faidgue Tests The determination of fatigue lives of components is discussed in Chapter 4, AMCP 706-201. The requirements for fatigue testing of critical components are given in Chapter 7, AMCP 706-203. The parts to be tested, the number of saniples, and the method of loading shall be specified. The samples arc cycled to failure, or for a prescribed number of cycles at several stress levels. The results usually are plotted as an S-N diagram. Basic information on the need, methods, and interpretation of fatigue testing in the helicopter industry is presented in Refs. 18, 19, and 20. By use of the fatigue test data and the frequencyof-ocurrence spectrum, the fatigue lives of the critical parts shall be determined by a method acceptable to the procuring activity (Chapter 4, AMCP 706201). If one or more parts have lives shorter than required, the flight envelope may be restricted or the parts may have to be redesigned and requalified.
Failure to comply with the fatigue-life requirement of the helicopter system specification usually will result in a penalty being applied to the contractor.
} \
5-4.1AMPLIFICATION AND NATURAL
tion prolAwm.• Such p'oints are uwd for test verifica-
FREQUENCIES A summary of thi field of rotor vibrations is conrained in Ref. 21, Although highly mathematical, this is a valuable reference for those seeking both general and specific information on rotor loads and vibration. The paragraphs that follow discuss specific anplification and natural ftequmicy Information. Fig. 5-14 presents natura frequency Information for a t wo-bladed, teetaing rotor, but the method of presentation and the design information are general enough to warrant a detailed discussion. Two graphs are made - one labeled collective mode and the other cyclic mode. These names stem from the types of forces and motions caused by the collective and cyclic controls. For the two-baded rotor, collective modes arm excited by even harmonic airloads while cyclic modes are excited by odd harmonics. The ordinates are natural frequencies; the abscissas are rotor speeds; the vertical lines mark the normal harmonic rotor speed range, and the radial lines Mre excitation lines. Coupled and uncoupled natural&4quencies are indicated by cutves with and witiout symbhols. The appnhc-be moi1e sha8r1 arem "-o~Wishown schematically. Fiapwise is nomnal to tile plane of rotation; inplane is in !he plane of rotation. Uncoupled means without blade torsion and feathering; coupled modes include these degrees ef freedom. The collective plot contains even harmonic excitatron lines, while the cyclic plot contains odd harmonic lines,
tion of 0e!'i calculated natural flreqncies. Without such verification to confirm or adjust tlhe fan plats during the first ground runs of a new rotor, several variations of tuning weights might be required to aw tablish correct trends at operating rpm, resulting in ccnsiderable cost and loss of time. V,e main objective of calculating rotor natural froquencies is to prevent coincidences (nronncem) such as Point A from occurring in or near the normal operating rotor speed band. Such steady-state rtoonances are not catastrophic, but often produce strell high enough to reduce fatigue life significantly, as well as fuselage vibrations that may require restriction of the flight envelope for comfortable operation. Generally, resonant amplification factors cannot be computed with sufficient accuracy for design purposx&. Once amplification factors have been determined for an existing rotor. fairly accurate predictions of loads can be made for vAriations of perameters; however, extrapolating these values to a new rotor involves considerable risk. The damping factors rOiur-iuR1 VIompuiing prturamanAr IN n hypi'M uaH IIfLC empirical and the value of the predicted loads necessarily is low. Progress in evaluating damping mechanisms has had to wait for improvements in both dynamic and aerodynamic computing methods because an observed level of response at resonance may be due either to a low level of force or to a high level of damping. Some'damping concepts associated with the fan plots of Fig. 5-14 are: I. All modes contain structural damping on the order of 0.50-1.00% of critl, which is relatively inof. 0.5-100 of•a• cri,,ti.i-iI. wh~tc n_*k.tk; significant. 2. Flapwise modes such " the ones shown in the collective mode plot ame strongly damped aerodynamically, while inplane modes are not. 3. Modes with strong intermodal coupling, such as those shown in th, cyclic mode plot, have fmequcncies that vary significantly with blade pitch; thus, they benefit from an effective damping mehhanism referred to as cyclic detuning. The cyclic
The circ!zd numbers identify three collective and four cyclic modes. All natural frequencies increase with rpm, but the flapwise frequencies increase much Ire~~~~tarde .... h._ . ...... . ..... ... £... . -h -. . . cyclic mode) because the centrifugal stiffness is a larger percentuge of the total stiffness in the flapwise direction. The collective modes show very little intermodal coupling with twist and pitch, while the cyclic modes show strong coupling, These fan plots ar' the primary de•ign guide for rotor dynamics. Myklestad's development (Ref. 16) made it possible to compute the variation of the unoupled frequences with rpm. Finally, the method of variation in pitch limits the resonance to a few Ref. 17 made it possible to compute the intermodal degrees of azimuth, preventing steady-state resonant oupling due to twist and ollective pitch. Accurate~ amplification. This damping source is at least as significant as the flapwise aerodynamic damping. computation is necessary because of the fine tuning required to avoid resonance throigh the eighth harmonic of excitation, especially for compound and 54.3 GROUND RESONANCE composite aircraft rotors having wide ranges of operating rpm. The helicopter shaft be free of mechanical instaTransient resonances such as Points A, B, and C in bility at all rotor speeds and opernting conditions the collecive mode plot (Fig. 5-14) cannot be (takeoff, flight, landing, ann taxi) regardless of the alvoided, but they cause no significant load or vibratype of landing gear; under the entire range of gross
A.i,'
. -K
T ,
"'.
•
-'
*1
0 COUPLED FLAPWISE NATURAL FREQUENCY < COUPLED INPLANE NATURAL FREQUENCY "-UNCOUPLED NATURAL FREQUENCY 3RD FLAPWISE SYMMETRIC MODE 1_600
2ND FLAPWISE SYMMETRIC MODE -
-1200
"-
S800
IST FLAPWISE
SYMMETRIC MODE
,-
400" r1UtYIAL urtKIA ~IUM1nGE
C.,
o~00
0
-
200
400
300
ROTQOR SPEED, rpm
,cy, L"
100
2N6FLAPWISE ASYMMETRIC MODE ) 1600,
/
1ST FLAPWISE ASYMMETRIC MODE /J
j•y
-
uj1200-,.
1ST INPLANE
~1200
k/o
SYMMETRIC MODE
,800 ...
:•-•
• .- ,
••
"- NORMAL OPERATING RANGE
300 200 ROTOR SPEED, rpm
100 "0
400
V. m 5-14. Typ"l Pb of Rotw Nafbdl Frq•eouy
Ovs ratiq Sped
-u
AMCP 71*302 0
A. STEADY-FOMa FMONANCIE6 i:l C SHAFT CffITILAL RWECXNCY (COULED.) * D WHAFT CRVITICAL IFFJW4C
MC NIL INSTABILITY
,
0
A.B.C - STEAY-FORCE FIESENAN(,ES
0
D. H - SHAFT CRITICAL FREOLIENCIES (COLIP.fUM E. I - %HAFT CRITICAL FREOUENCIES (UNCOUMD1E) CC. HI - SHAFT CRITCAL INSTABILITY REGIONS
(UNCLED)
NE-P~FIVV EXCITATION OATING S"srw DIN REGIONIN
~
~
JIK MECHANICAL. INSTABILITY
LM
REGIONS
ONE-PEAlI4v EXCITATION ROTATING SYSTEM
ED TR
PYLON
SPLER
do
Coe
)
I
a PltR
-f-t er weight ancodtos hogotteeteetn Ge5-5 z yicciuCH.AN MEaucrncaeifc jICnALv u' ~ ~Fi ~ PYOseisoftssisseilc yCape ,AMCPI - heac 203.tets canhc relaed, orEliIONatd rvidedT thatýCA itcnb eosrtdt h aifato oftercrigatiiythtte eicpe i re f mehnia intbiiy Ocurec of' the intblt 1rbe0 sulyi of /ehla rmr l alr ~ h yTem d' in information isRAIN cotie inKf.2,2,ad2.Te
u, m
MASTi
2I
P
UI
YLO
iue51.Todge-f-fed0
Ck
.Po
sytm uhgah Clma lt)peetbt L6 feu~ n tblt nomto codn oc etosdti nRf 2 Pteta iStaiies-Athog F-idiae M h rqunypo i i.5-3accasde1i h' aegre nacrac wt oeo xi ain I ot n :odir eoacs(eo
Teihe aonalyitos,ofd thermuchanica ithbilityeof two2. Poits C-6andere-om DsafcrtCalta spes eci beaieds hingedt roos dpiffers fro thapte of rotor W0th sytem.byc gro~t-aphsc fCole)a pltha) oresat bthe 20.Threse oremosbladsIn therlxd originalinalteis (rof. creupedc (Pandtblty C)formtdonuncoured(Pint suport to 22)itd a sued that emosupprate thcab syteobeo the s ytsato eoste fdetaciledin Rf.2 rotoranincaludingtailthe pylon lpndin in. frndncand F:caececplonansfid i. Fig.t5-1 blad gecuren oulbhe restablvdintypobasnlem sal isr-f a tr maagrkein bcoundanies oftthe moechanexica freedmsuurtite a rmr of alr ntwo sysem.Additia(oreintabiliyron(sl-:iedmdsde ongtv andfaftaandnlateral)adirectionsfsTh22,n2cessntate. the dami. forcs. Theand8 porintsalwayrs oacur atreoto dicmutation othatfllw suffeti ent mass andaeletivel speed greter sthady thre)f~thi ooat ton-alf urcr thefr eahmdofviewgvr nCaption of the0-21 ItsTeaaysiem support. ouplednatinaludrequetwo degre of) acht cudirection.Th A)te ind mode a(reoftesPpornt sse support eachM111 dD tha futhr Asu. be90.1 treated ine M edetlyl Mv 1O rVL yot ationalsytmfrequency(Poit geeatdb)ti The analysis osfe o the opt cehnclistblt ftoupe anly.Pins Cie ind Fi.D: shaft-cr .I aditicaonals(eci blntualed fhingedotosoiffesfo moel th whichohs at wlmets steady forc res-on-ani anadtioal fo, e ochraftCrthea ofbthwa 22)' asumedd rthatineupor systems t eus, the srysrtnem H-I, and two adtoa ntbiiyrne
\
uniscnb folottedaf and aclateda prnsn
rtrthe in ei
for achmodeof he sppot. ibraionof I wa
andnL-M.Pon G) Th anaysi of
e 1;1 nclues wo dgres2o
ii-NORMAL OPERATING SPEEDK -NATURAL FRLCKJJCY
NORMAL OPERATING rb RANGE 0 200.9
FREQUENCY OF FIRST INPLANE SYMMETRIC 3ENDING MODE
20 0
SMINIMUM 1.b
O
ROTOR
IN J
MINIMUM FREQ >: 1.ROT04_ 151l.0 1
/• ",
K
STEADY FORCE RESONANCE
BLADE MAST7
•
T
1ST MECHANICAL INSTABILITY
8 1T SAFTLOWER CRTCi ''
•0.5
o
DE
0.5
1.0
)
1. 11500"
I
I&oI A
z
BOUNDARY
II
MAST-
1.5
MA
2.0
HI
25
NORMALIZED ROTOR SPEED _aL flo Figure 5-17.
5-43i. II
Two-degree-of-freedom Coleman Plot Sho*ing Sati•faction of Minimum Frequency Criteria for Two-blade Hingeless Rotor
Two-bladed Rotors Without Hing-s
ran h. ahn.,=
thai t*wn...4aI
Aa i..ia..mnenr-
r-rnntu.A
the coalescence of tle two support system fren,.-n.ac
nni*
I
,.•tmni,.ultI
l
nin
tha
can be approximated by rotors with effectivc hinges; thus, the analyses discussed in par. 54.3.1 also can be applied to this case. Because of the location of the rotor natural frequencies, the ground resonance requirements can be me! by satisfying the following criteria (see Fig. 5-17): I. Sufficient damping must be provided by the landing gear and undercarriage structure of the airframe to remove the instabilities (such as D-E) associated with the rigid-body cegrees of freedom of the airframe on its landing izar, The exact amount of damping rquired to stabilize the region D-E is not knowi,, but it is probably Ins tharr 3% critical. For configurations with skid gear and pylon isolation, the inherent damping of the system always is sufficient to prevent instability,
operating range. 3. The first inplane natural frequency of the rotor at high collect~ve pitch shall noi be less than 1.15 times the rotor speed. This assures that the mechanical instability associated with the coalescence of the rotor natural frequcncy and the lowest support system mode (Point C) will be above the overspeod operating range (110% rotor spceW), and prevents excassiva response to steady-state one-per-rev excitation (fl - 11 line).
2. The mast frequency in the rotating system (Point B) shall be no less than 1.2 per rev. This guarantees that the mechanical instability associated with
tion of motion associated with out of-bvlancc forces; the stability ranges given in Fig. 5-16 by D.E and H-I are reduced to s'mple resonances of the support
5-22
54.3.3 Multibladed Rotors The analyses discussed in pars. 5-4.3.1 and 5-4.3.2 are special cases of the classical analysis given by Coleman (Ref. 25) for mul'ibladed rotora. For rotors with three or more blades, there is no preferred direc-
A•P 706-202 modes with the operating speed of the rotor. Thus,
hcli.-opter and compound; and under each to discuu
the number of degrees of freedom is reduced, simplifying the equations of motion However, the analyses of the mechanical instability ranges J-K and L-M in Fig. 5-16 are identical to the two-bladed case and must be treated with the same considerations.
fixed and rotating system divergence and flutter, or static and oscillatory aeromochanical instabilities.
5-4.4 FLUTTER ASSESSMENT
Historically, relatively 1i'tle attention has been given to fixed-system divergence and flutter for helicoptcrs, as other design end operational requirements such as static strength, fatigue life, and operating speed havc precluded divergence and flutter.
The other types of potential rotor instability arc divergencc and flutter. The discussion that follows supplements the analytical review given in Chapter 5, AMCP 706-201. When the developing acrodynamic force simply overpower the elastic constraints and the motion exceeds some presdlected bounds, i.e., goes unstable, divergence has been reached. Flutter usually involves 2 change in and coalhacence of two or more system natural frequencies because of dynamic or aerodynemic effects, and the coupling of oscillatory motion of the lifting surface with the airstream in such a way as to derive energy irom the airtream to increase the motion. The first formulation of the flutter problem was published in 1934 and subsequently republished as Rcf.a. 26. Each potential fluster problem is related to modal couplings that are configuration-oricnted. and the
5-4.4.2.1 5-4..2.11
Helicopter Ixcd System
5-4.4.2.1.2 Rotating System A number of rotor acromechanical instability problems were encountered in the period before 1960. Solutions usually were worked out by trial and error long before they were understood mathematically. Included were problems such as weaving, pitch-flap. pitch-lag, and pitch-conc instabilities, and stall and binary flutter. Of thesc only o11c of recent occurrClece is stall flutter, which has been actively restarched (Refts. 30 &iid 31). This work has provided an understanding of the problem and the ability to predict the stall iuwitr houndary with reasonable accuracy. The basic dexign changes that sol~ed most of these problems were overbalanced blades, torsionally stiff
number of such couplings is very large. Furtier, in
H-lades, and, with the advent of hydraulic boost, in-
each specific configuration, the stability equations involve a large numbcr of parameters whose meaning and measure are only made clear by a rather precise analytical diagram or model. Thus, a specific, rather than gen-ral, method has evolved. No method has yet bearn devised for writing mcaningful specifications and simple instructions for designers for the preventior' or avoidance of these instabilities. Notable attempts toward simpiification are given in Ref. 27 for fixed-wing aircraft and Ref. 28 for helicopters, and a recent attempt at ordering and classifying is given in Ref. 29.
creased control-system stiffness. During the design of conveuit.onal rotors, the current practice is to forego elaborate calculations. The only mandatory check for main rotors is that the chordwise location of the effective CG of the blade be forward of 25% chord, and preferably forward of 24% chore- For unconventional designs, Serious considertion should be given to detailed quantitative analysis. The list of known problems should be checked to see if an analytical solution is available, preferably a method that has been checked against experimental results.
5:-4.4.1
5-4,4f2.2 Compound
Carrels Crlittr static and oscilla-
"7he most comprehensive list of tory acroclastic instabilities compiled to date for hell-
coptem is presented in Ref. 29. Several specific con•3-figuration-oriented problems are named, and' formulas are given for determining static and oscillatory stability boundaries. This list could be extended to form the basis for a usable specification dealing with the aeroelastic phenomena basic to helicopters.
54.41
Dedpa Cosu.derations
To discuss design considerations, it is convenient "toclassify the applicable aircraft configurations as
54-4..2.2.1
Fixtcd System
In the development of compound helicopters, considcrable a:tention has been given to divergenmc and flutter due to the extension of the speed range beyonw that of conventional helicopters. Conventional prvctices, such as those outlined in Chapter 5. AMC-706-201, are adequate. In one known case, signim.!cant buffeting of the vertical fin was encountered due
to the impingement of disturbed air from the hub and pylon. The problem was solved by cleaning up the flow.
5-23
5.44.2.3. beaft Sydown A grmit deal of rowesh has 6w. accomwplihdW during the ertasion of the helicopter speed range by compounding. Ref. 32 is a Sood auniniary of the early work in this are. This work showed that increasing advance ratios and blade tip Mach numburs require prob ruive unloading of the rotor and reduction of rotational speed - e-v= down W zero. The associated dynamic phenomena arvi continuous and trackable until umro rotor speed is approached. At high advance ratios, thrust and flapping acirrol are difficult because of high sensitivity to gusts. The Principal dynamic probkv is limit-cycle flapping instability. Ths instability produces both harmonic and nonharmonic flappiri, the latter being visible as a weaving c! the tip path plane. The history of this probiemn is sketched and the picture clarifiled in Ref. 33, which shows two azimuthal regions of inaitability, one on the advawacing side due to negative spring rate, -and one on the retreating side due to negtative dahipins. Measures for stab~Iinga both are dimsausd.
GIUST LOADINGS The need for mea~ninful Sust loading speedrictions has increasd with the development of high-performance helicopter and compounds. The first dofinitive work was Ref. 33, in which the nature of the problem wus elucidated and some solutions were offered. A morm precs tmctment required more sophisticatid analytical methods. One such method (an extecnion of the rotorcraft flight simulation niethod discussed in par. 5.4.1.1) was developed and ueed for an extensive study of the problem (Rd. 34). This refernceK also reviews the state of the art. 7be problam of helicopter response to gusts and acmre of the design consideraticns air discusd in pars. 5-4.6.1 44d 3-4.6.2 that follow. 5-4
S4.6.1
Dh UUofS@ t'iot G"~ Problem
Rotary-win& aifrcrft expcrrince milde reactions to gusts than do most fixed-wing aircraft. One of the carliost reports of this difference presents qualitativo itutctions of two pilots on a dual flight, ont"n a helicopter with side, by-sick rotors and the ( cr hia a S-dg ArnIJT'rnr 7.nAnvmac rixed-wing rimlane A similar test wsir mincmwf Ioate by NACA with instrumenwaion to measure norConventionial rotcrs, bucause of t"er relatively low mlfraibohtypes of nircrsft flyitag through turdisk loadings, experience negigible acoustic loading. Airflow over the blad!e aurfacca; cani reach somic veun ar The relerively mild reaction of the rotary-wing airlocity loalWly and momientarily produce a shock ~ Aeodynmicloadng vriaions~ ~ craft is not substantiated by the simplo theoretical exmanly Atherodyam ic, whaich geariationthe t phriprzaaions currently in use, p cartclarly those that istic acoustical sigrature referred to as blade slap cvovcfrmixdwneprca.Fg.-8sos an example of gus. lo-d factors resulting from .AurpRotors have operated in such an environmeint for edged gusts, computed ey the pzocedure in Ref. 35. yaswith no evider. x of structural fatigue caus-d b eithr r cutclain.BaesabyoyaiThis procedure it conservative in that it neIects stall eihasen showni tor bedrag-eloatdng henad, slap and compressibility elfects and assumas imsantaneCsacustic
ioadiing occurs iapiane along the axas of
.-.
I. - .. .4_--.~
.. ,,
an..
... A."..
I.L~.
,ping. ty, atid black which thooretica. gust load factor may be reduced. w I--) "vdsags leito atrb
maxmum evveng'h of the blade skin. Proprotors itnd propellers located in closn proximity to either a fustlag or a wing can experience hkome degree of acoustic loading. However, the rotor sir presuurc impingement on W,~ ftuaselgor wing etruc* turu in an order of magnitude more si~gificmnt thant4. acoustic leading of the rotating system. 2. Tail rotors located clo, to a taIl boom e~parionot aeroclynamnic kw~ming catuWe by the pirtial blockage or airflow an" the intuf~rx.n% by the fin with Wk a)
blade suarfac prooswc field. As with amnvetional
-
___ii
1
1
NO COM.iSSIBIIL ITY OR STALL 50 lps SUDDEN GUIST
II-__
main rotors. there is no evidenc of measurable rcouak~ loadling. Ductnd fans used as antitorque devicest, because of their relatively high disk loais-rgs, 01 0.2 0.3 0.4 may pitodi*c rrneter acoustc lo~adi than con-ADAC AIp AI DAC ventional tlil rotors. However, the duct am rounding the fan experiences highar loadings than do the fan Flk,'Me 5-18. Gum L*W Factor Cosapaed for- the bladck UH-t8 Heliopter Uslg Umar 7%moy
7"620
_____________AMCP
This Suit alleviation factor AKin given as a func~tion of rotor disk loading, and is equal to itnity for disk loadvils gieater than 6 lb/ft, as shown in Fig, 5-19. At high speeds and for disk loadings greattr than 5, load factors computed in accordance with MIL8-9698, inr'ludi In'S the gust villeviation factor, aec very high. When manicuver loads are surmrimposed on these gust loads, an uiirealistic situation results. On the other ho'nd, studies indicate that the thrust o, - rotor actually decieases with increasing advance ratio. The aerodynamic limit shown in Fig. 5-20 is calculated by a digital method based on Ref. 36 which includes the effects of stall aaid ompressibility. Also shown in Fig. 5-20 is a practical limit for thc same rotor based on flight test data. The practical limit is a result of oscilinoury rotor loads and stall flutter effects, and iý the controlling limit on rotor thrust capability at high advance ratios. This conclusion is supported by Ref. 30. Unloading the rotc'r by adding a wing would give the rotor a gr eater margin to accept gusts. The ad vantage, however, is not as great as might be expectei, because the rotor usually will assume the larger share of the lift increase resulting from gusts, as shown by
*gust
-'capability
i
The gust alleviation factor given by MIL-S8698 was found to be unrealistically low for a rotoi unloaded in this manner. An early attempt at treatncnt of gust effects on rotary-wing aircraft is reported in Ref. 33. Sinesquared &,ust shapes were considered instead c-f sharp-edged gusts, and a mass ratio comparable to that used in fixed-wing anaiysis replaced disk loading in the determination of the gust alleviation factor. Fig. 5-21 shc ws the results of that study. It suggests a gust alleviation factor considerably smaller than that * given in MiL-S-8698. AThe scope of the study, however, was insufficient to define requirements for all types of rotary-wing aircraft. Furthermore, gradual penetration~ into the gusts, nonsteady aerodynamics,
orne aeroelastic feedback wert not considered. Subsequent studies (Refs. 34 and 37) attem~pted to remedy these deficiencies and to include onstcady aerodynamics as well as 2dditional variables'such as gust shape daid intensity, forward zpeed, dis.'Aloading, thrust coefficient-solidity ratio, and advancing tip Mach number. The design considerations derived from these studies arc reviewed in par. 5.4.6.2. 5.4.6.2 Gust Design Considerations Fo h tde r es 4ad3 eea rn cipromnthestuiens ian befs 34and 3he gstleveraltpionfctponclusgioens can be drawn. hregust allservation. fahtorseKfgiroerin M s rat-io tre too contervaine.Kb Th usrfartrms ai ,t eemn analogy with the fix~ed-wing approach, as suggested in Ref. 33, also does Aot give satisfactory results. Never theless. gust loads cannot be ignored in rotor design. Pending preparation of criteria to replace the gust load requirements of MIL-S-8608. alecinative methods for dettrimnation of gust load factors may be used, subject to thc prior approval of the procuring activity. ,b
ICAI ALV~fl(NAVI1C 114[O111
-/copter.
-
!AXMIV.v
HUOCBFjA.
______
Ify
UiAAAE ýi 0.10 -1 ,-
MP.UVERING
-
I~T _0
/
iIAýCIICA[ LIN'I1
-
0
02
04
It
08
ADVANCERALID
RurLmissaFnconl
adsc fim52.Bi
~~~0.6I-4,
iotono RLitsas
docRai
-~I
~SIVj,,
o
0.
--
VISK
4 5 LOADING, pst
321
-
'
1123
HEWtLTS.KF.[
V~ '4
NIL-S-89
=10.2
-)
-7=
I"
7--- 1 n I -------
008
*
k
r0
:'
4
15
8
6
~~Flgure 5-19. Gust-sillevlation Factor (MIL-S 48698,
CUMSSAI,
Filgure 5-7 1. Results ot a Lead Gust Stu.&;, Conapaied With Military Specification Requirements 5-25
For. all the helicoptersgrid compounds investigated ina Re('. 34, the rotor gust-load ratio AT/T~w7 can be expresse by the empirical expression AT 0.57 085L ~bility + C (5-7) -
-
.
T~,
(r~
~later
wheac\ 'ae C - empirical constant, dimensionless 11 - wing lift. lb LE/T,..M - wing lift ratio prior to gust enccunoter, dimensionless This relationship givet reasonable accuracy with appropriate conhervacism with C - 0.2 for semirigid (teetering) rotors and C ft 0.1 for rigid and articulated rotors. For a compo.und helicopter, the wing gust load should be determined by conventional fixedwing methods. However, an additional a~lleviation of the wing gust load, owing to the interaction with the -rotor, was found to be related to the rotor thrust coefricient.-solidity ratio CT/ar.'With further refinement, this approach may provide tin acceptable basis for gust design requirements. The relative affocts of various parameters on gusst response amc summarized in Table 5-1.
m~m
5-6.7 TORSIONAL STABILITY ircaftiror arcrft agies irs-deeloed Gasurbne Gasturinee~gne fist eveope fr I~ eluded fixed-turbine engines for both turbojet and turboshaft configurations, and free-tutbine engines for driving rigid loads oir loads with natural frmquenc"e far above the maximum response range of the engine governor, Sophisticated hydromechanical TABLE &I1. THE RELATIVE EFFECTS OF
VARIOUS PARAMETERS ON GUST RESPONSE
governors were optimized for thwc sy ns;uhowever, thesw governors proved unsatisfkacory for bellcopter applications. A serious drive systet instawas predicted for the XH .40, the first helicopter using the firee-turbine e~ngi (this analysis was publishvid in Ref. 38). The prddficWe instability occurred as predicted, but the consequences were not serious. Means to stabilize the syttemn had 'ace. provided and thMe test program was able to proceed, alihough with a rather sluggish governor. ExtIensivo analog oomputer studies have shown that serious penalties would be i,-curred if the helicopter drive system were., mo'.ified to solve the prob1cm, but that modificatioa of the governor reaulted in only a slight pona'ty. A. a result, Ref. 39 was formulatoz. This piubfication ansigns responsibilidics to both the engine and helicopl-ýr manufacturers in order to insure early rmog~nition and solution of the drive system stability problem in future applications. The discussion that follows suppitleents the reviews rif drive system torsional Ftmbfifty in Chapter 5, and ol the enigine/airframe integrated control system in Chapter 8, both in AMCP 706-131. Discussoo of Problem 5-.. Thc dynamic characteristics of three systems are involved: the gas producer and its govcrnor. the pwrtrieadisgvroadtehlcpe drive syslim with one or more low natural freqece.Tesse qain fmto oiei qucisThsytmeaiosfmtinovnently can be put in transfer function form and atrayed as a block diagram (Refs. 38 and 40). During throttle movements, tht. gas producer governor controls the engin. At steady-stlate throttle conditions,
the
______________speed
PAAEWEFFECT DIKLOADING ROTO 19-D-i SIIDIT't RATIO. C,'/
MAJOR EFFECT(SEEEQUATION 5-7
urops bc'ow the selected value, additional fuel flow is called for to nullify the error. Accordingly, when tht power turbine overspeeds, fuel flow is decreased. This cycle of events can be stable or unstable depending
DERAkE EFFECTAl HIGH VALIJES OF C /. DUE TO LIFT SHARING %ITH A r,;Ný'The
upon system parameters.
ADOTYarPE
SOMEEFCT. DEPENDoS ON OYNAMILS
NMBIER OFBLADES
LITTLE EFFECT
NUIMMER OF RO~TORS
INCREASED EFFECT FORTANDEM
usually can be simplified to a system consisting of the main-rotor and power-turbine inertias connected by tihe effective shaft stiffness between them. This is a
CONFIGURATION-
single-degree~-of-fieedom system in which the main
*CONSI COmPVuNOiNG
I
LIfll.E INFLUENCE
powar-turbine governor controls Dower-turbine by modulating fuel flow. When this speed
FRADVELOCITY AND
ADVANCING-TIP MACH
KM.3IS LOCK NUMBER ______
tia is much larger than that of the power turbine. This
~~~~S.IGml ~LOAD WITH INCREASED OC
_____NLUs&B
?It!C4 FLAI COULIUNG PITCm CONL COERLING 1OWISGH11 IN COLL.ECTIVE S%~ 9ETB
5-26
rotor can be considered as nodalized because its iner-
LITTLE INFLUENCE. _______
LOCK
____
LITTLE
~FC
AFPRECIASLE EFFECT APPRECIABLE EFFECT __________
drive system for rotors without drag hinges
____
modeTO usull hGUSaurlfeTec eow5Hi eow5Hi maeuull0C1atrlfeqec lightly damped, and is continuously excited at a low level by rotor control m'otions and external transiResulting oscillations in the power turbine are sensed by the flyball governor, which modulates the fuel flow accordingly. With a governor
Jspeed .~ents.
i° •:Ž•.ig:ss. .. •.:
:•L•.•L•, "7••-, • 2
.
••,g....zL*•.••,
optimized for a systcm with a high natural frequency, the oscillation of the torque at the turbine wheel that follows will be so phased that it reinforces the original, lkw-frequency, drive system oscillation, Unstable torsional motion results. It is not feaible to stabilize the system either by inc:uding mechanical dampers in the drive shaft or by stiffening the shafting sufficierntly to move the natural frequency out of the response range of the governor. Furthermore, the gains and time constants of the engine and governor can be varied only within narrow limits One effective solution is to use a small amount of valve oyerlap, which allows the iiyball governor to oscillate the fuel valve a small amount without modulating fuel flow.
the respective hinges. The lag dampers (shock absorbers) prevent unstable blade oscillations about the lag hinges. Dampers usually are not required for flap hinges because of the amount of aerodynamic damping provided by flapwis Jade motion. The design considerations for a typical articulated rotor, the motions and loads for each hinge, find thtir effect on the helicopter are discussed in par. 5-5.1. 1.1. This rotor has the flap hinge inboard, then the lag hinge, and then the pitch axis hinge outboard. The effects of reversing the hinge arrangement are considered in par. 5-5.1.1.2.
54.7.1 Design Coelderations Ref. 39 establistes an effective, three-phase procedure that adequately deals with the problem. Briefly, the steps include: 1. The engine .doaigncr provides as much flxi-
Fig. 5-22. The hinge arrangement of this rotor also is dorscy S-si,the Brintoi Model 171, typical of thte the Alouette H, ano the Russian Mil 6, 8. and 10 aircraft, among others. The hap, lag, and pitch hinges on the Model 107 TrQtor have oillub icqtcd, cylindrical roller bearings. The three flap hinge have a common, centrally lIcated reservoir, while eazh of the other hinges has
bility as possible in-the engine governor parameters. 2. Early in the preliminary design, the engine and heliconter denioners evechangp syuem rnie.rnodrm
nnl
each designer conu4cts an analysis of the system. 3. The helicopter designer selects the Optimum Seveoal efhicient e minputer methods of anyysis now exist, including antlog, digital, and hybrid. The analysis can be complete, including the nonlitioar parameters for the full range of enginq ptperation, or a perturbation analysis can be informed in which an operating point is selected about which oscillatory stability is detetmined. The complete analysis is much S~small more complex, but it determincs transient rcsponse n.Ad droop
ass ... a.I as ............. ......
sa.l.end#.~j.
-ip..
L.
tB "" •BatVB
Ca"
5-5.111 Typical Articulated Rotor Coesideratloas rhe Boeing-Vertol Model 107 rotor is shown in
t own reservoir. Each reervoir one or more sight glasses to indite oil level. has Radial, positivecontact seals retain the lubricant in the bearings. The flap hinge is .Offset both radially and in the direction of rotation. The radial offset of the flap hingv axis is small, approximately 1.7% of the blade radius, and is as close to the rotor cnter as shaft and hinge sizes permit. Boqatts of the small radiatl offset of the flap hinge, the gwntrol forces generated by this rotor come primarily from thrust vecror tilt, with a small contribution from the the vertical veca component cl nt ofa contribution from the flap hinge forcn isee Fig. 5-2 (B)). The offs t of the flap hinge in the direction of rota-
lysis u1sually i- aequate for dletemining stability alone.
tion (also known as the torque offset, and shown as dimension "a" in the plan view, Fig. 5-22) is chosen
5-5 BLADE RETENTION
to satisfy two requirements: 5. To equalize loads on the flap-hinge bearings when the blade is in the lag position, corresponding to normal flight torque 2. To avoid reversing axis, motion of the flap hinge bearings due to blade lead-lag motion in normrtal flight. The flap hinge cylindrical roller bearings withstand blade centrifuga. force, alternating loads due to blade lag oscillation, and blade vertical shear forces, and experience one-per-rev flapping oscillations of :1:4-6 deg. Thrust loads are carried by a b;onze thrust bearing. Permanent stops prevent excessive blade droop or flap motion due to winds while the helicopter is parked or during rotor shutdown. These stops are set
5-5.1 "CONSIDERATIONS RETENTION SYSTEM DESIGN The fully articulated, gimbaled (teetering), and rigid (hingeless) rotors are Jescribed in par. 5-3. The blade retention requirements for each are different, and are discussed in the paragraphs that fclow.
-_
.
5-5.1.1 Articulated Rotors The fully-articulated rotor system provides freedoa of Mlade movement about ilap and lag hinges in ")response to aerodynamic forces resulting from pitch change and/or flight conditions. This flap and lag freedom reduces the flap and IA; moments to zero at
5-27
AMCP !L&.202
.I
*thus
3
so that no contact occurs in normal flight. The flap stops pre-vent prissible ove-tumning of the blade; droop stops prevent blade/fuseleg contact. Slade angular displacement about the lag hinge stop to stop -varies inversely with the radial position of the hinge. However, larg angular lag displacement adversly affects: 1. Blad%-,toý-blade clerance in a tandem heclicaptar 2. Pitch arm kinematic error 3. Lag damper stroke. The Model 107 lag hinge is located as far iaibomird as possible, consistent withi the lag displacement considerations listed, in order to keep the mans of the pit -hi hinge and the blade retention joint inboard and reduce the centrifugal loads. The lag damper, in addition to meeting requiremeats for stability of the lag motion, functions as the lead and lag stops of the blade. Design loads for the stops are the rotor starting condition and the rotor braking loads at shutdown. The damper is positioned so that the centrifugai load on the damper will be along the piston-rod axis and will not result in pistonrod bendialz or internal-bearing wear. Lag damper end Lem ririgi art lin bd with Teflvin fabric for ar~airite-
The rolling element beazings of the lag hinge are subject to :cntrifugal loads plus reactions from blade alternatint and steady bending moments. Vertical shear loads are carried by bronze thrust bearings. Blade lag oscillatio: is in the order of * I dtg. The cylindrical roller bearings of the pitch (feather) hinge react blade steady and alternating bending moments and shears at the root of the blade. Typical osciliation angles in forward flight are *4" deg. Tension-torsion straps, consisting of many slotted stainless steel elements, react the blade centrifugal loads. These straps twist easily, providing freedom for blade pitch change with negligible effect on the pitch control linkage force. The pitch arm connects to the upper end of the pitch link nearly in line with the flap hinge axis; the locations of this attachment and the lower attachment of the pitch link to the swashplate were chusen to obtain favorable coupling of blade pitch with flap and lag displacements. Two taper pins in a multiple cievis joint attach the blade to the outboard end of the pitch housing. The dclvis provides freedom for manual folding of the blades about or~e or the other taper pin, depending upon the required direction of fold.
nance-froe operation.PICAR
PITCH HIOUSINdG
A0ORHU
I
FLAP H4INGE RESERVOIR
SIGHT GAGEPICSHF
ROTOR
FLAP
I
SHAFT
HINOF
LAG
Figure 5-22. 5-28
Articulated Rotor (Boeing MoMe 107)
ROTOR SLADE
_AMCP M06202
)4 Ieirersed Hine Articulatiom 5S=.I.I coincident lag and flap hinge arrangement The shown shematically in Fig. -4 is ued on all production Sikorsky helicopters from the S-55 through the S-65. •hc radial location of these coincident hinges is ro~hly 5%of the blade radius. Good controi power results, permitting liberal CO travel in these single-rotor helicopters. The loads and motions resulting from this hinge arrangement and the retention methods u~ in a typical coincident-hinge rotor (Sikorsky S-61) aliscussed in the paragraphs that follow, As the S-61 lag hinge bearings are mounted in the star-shaped hub, normal blade coning and cyclic flapping result in sizable vertical thrust loads along the lag hinge axis. A pair of conical roller bcarings transfers both radial and vertical thrust loads of the lag hinge to the tApper plate of the hub, while a cylindrical roller bearing transfers radial loads to the lower plate. With this coincident-hinge arrangement, the flap hinge leads and lags with the blade. Loads on the lead and lag bearings of the flap hinge are equalized by a :...... ica!
.
betho Ion Inao. oLic Avinl Io,,n*;,•A
thrust along the flap hinge axis is due to blade chordwise shear forces. These forces are low and are carried by thrust faces. The pitch (feather) hinge isjust outboard of the coincident hinge, and blade moments and shears are of low magnitude, with the primary reaction for this hinge being the centrifugal force on the blade. A stack of angular contact ball bearings arranged in tandem carries the centrifugal forces in a very cornpact arrangement. A radial bearing pre-loads the angular contact sct and assists in carrying moments and shears. At least one helicopter (CH-47) has the pitch hinge inboard of the lag hinge. in this configuration the lag axis rotates with blade pitch changes and is perpendicular to the flap hinge axis at only one blade pitch. The weight of the rotor head is thus reduced, as the blade can be removed or folded at the lag hinge and an additional attachment joint is not required. Hinge loads and motions for this configuration are similar xo those of the typical rotor arrangement first discussed, with differences resulting mainly from the moments and shears at the different radial locations of the pitch and lag hinges. Control system loads do not differ significantly. This rotor configuration, with the pitch axis hinge inboard of the lag hinge, results in two major characteristics: 1. There is no kinematic coupling of pitch with lag. 2. The "parked" helicopter has reduced blade/fu-
claiga clearances in that the weight of the blade in a lead or lag position results in moment about the pitch hinge, causing the controls to "drift" and the blade to droop below the normal position. 5".1.2 Glmibakd a Teeterig Rotors The gimbal-mounted rotor and the two-bladed teetering, or see-maw, rotor have blade pitch-change hinges rigidly mounted to the central hub. This hub assembly in turn is free to pivot with respect to the mounting structure in response to one-per-rev blade forces, thus minimizing loads in the blade root and the hub due to first harmonic flapping. Coriolis forces in the lag direction similarly are reduced. Alleviation of these two types of load has a significant effect on rotor-hbud strength requirements ard therefore on the weight of the components. Elimination of the lag hinge and lag damper reduces maintenance requirements but at the expense of providing strength for lead-lag moments that do nut go to zero. Also, controllability is somewhat lower with these hubs because it results from thrust voctor tilt alone. Gimbal-mounted Hubo Two-bladed hdbs, fully gimbal-mounted on the rotor shaft, see-saw about one axis for cyclic flapping and arc tilted about an orthogonal axis for cyclic pitch control. Collective pitch is input by individual links to each blade. The OH-13 and the OH.23 are examples of lielicopters that use this hub. The gimbal pivot bearings react the rotor lift forces and also transmit the drive system torque. When the rotor plane is tilted, the Cardan joint characteristics of the gimbal cause oscillating speed. torque characteristics in the drive system that must be considered in
5-.1.2.1
t
i
-- -- ,: ..... thc.'Ctc,-. or..:..1 gf.4.
.....
bility in the gimbal hinges is required to react the inplane rotor forces. The pivot bearings on the gimbal axis parallel to the blade span escillate with cyclic feathering of the blades while the bearings on the axic normal to the blade span oscillate with flapping of the rotor. Thus, the bearings on the pitch axis are required to accornmodate only the collective pitch motions of the blades. The pitch axes are preconed to reduce the steady blade flap bending moments on these hinges and on the hub structure. The hub structure containing the hinges is underslung below the gimbal pivot so that the vertical location of the CO of the blade assembly in the normal flight position is close to that of the gimbal pivot point. This reduces the chordwise oscillotion of the blade. CO when pitch changes are n.ade without flapping, ar discussed in Ref, 41. 5-29
,
Moments and &hear fort for the pitch berings amrhighe than thoe of an equivalent fully articu1taed rotor. Motions are of the same order of magnitude but, as noted previously, do not Include tie ocillations due to cyclic pitch. Retention methods are the same as those previously described for the pitch axis of the fully articulated hub.
for and significas of then coupling effects upon the stability of the rotor system are discussed in detail in Chapter 5, AMCP 706-201. The advantage of the hingeless rotor "am indude the high level of control provided by the trafor of moments. The system may be physcally iniple,. but the strength required to transmit the forces and moments across the blade retention system can cause the retention to be heavier than is true for other types of rotors.
S-t1", Teetearla Huls Anothir common two-bladed rotor hub configuration, used on the OH-58A and the Fairchild-Hiller FH-l 100, is . hub free to teeter about an axis normal to both the blade span and thu rotor shaft for cyclic Capping. Control linkages to each blade change collective and cyclic pitch. This type of hub has preconed pitch axes underslung below the teetering hinge, and, in general, the loads and motions arc similar to those of the two-bladed, gimbal-mounted hub. However, larger chordwise moments are caused by the lack of full gimbal provisions, and the pitch axis bearings must accommodate the oscillatory motions of cyclic pitch as well as collecti,\ motions. &5.1.3 Ri5d Rotor
(~NSIDflATIONS The design considerations applicable to the use of rolling element bearings in rotor blade retention systems are reviewed in this paragraph. Additional discussion of bearings, for blade retentions and for other applications, is found in par. 16-3. Other components associated with the blade retention system that also are discussed in this paragraph are: lag dampers, lead and lag stops, droop and flap stops, and drvop and flap retainers. M5.2.1 Rolling Element Bearila
I
The "'rigid," or hingeless, rotor blade retention confiSuration attache the rotor blades firmly to the hub, which, in turn, is attached rigidly to the rotor mast. The blade retention system must be capable of transferring forces and moments in both flapwise and lead-lag directions. The blade centrifugal force may be reacted by any of the conventional methods (a tension-torsion strap, an clastomeric bearing, or a stack of antifriction bearings). The retention must provide "apitch change capability (both cyclic and collective), and the centrifugal force must be. reacted across the
*
pna
Rolling element bearings are widely used in rotor hinges. Experience with them has been good and the technology, which is based upon both analytical methods and empirical data, hat been verified by extensive service experience. Gene ally, these bearings are compact. Bearing friction as low and has a negligible effect on hub loads However, the effect of pitch axis bearing friction on control system loads should be evaluated; on large helicopters this friction usually is low compared with the aerodynamic and dynamic loads, but on smaller helicopters the effect
SNo
hinges are provided for flapwise or lead-lag motion and the only flap or lag movement of the blades relative to the fixed support structure is due to structural deflection. The specific layout of the hub and the retention determines the manner and extent to which these daflections couple with and affect the pitch motion of the blades. As with other systems that do not incorporate a flapping hinge, a precone angle usually is built into the blade retention for a hingeless rotor. This built-in angle helps to alleviate the flapping moment that the retention must react, The geometry of the retention also may include sweep and/or droop of the spanwise axis of the blade relative to the feathering hinge "axis. The direction and amount of these alignments, together with the specific geometry of the blade pitch control input, define the feedback coupling between blade u tion and pitch control input. The necessity
S.: 1
5-30
Sz_%'. . .-
...urn...m •,, ,-,,e-
.... ass
...aaaa-.
iliy tic
as n a~ifst.
The failure mode of osczllattng rolling element bearings most generally encountered is gradually progressive spalling that results in looseni-ss, heat generation, and aircraft vibration. These factors have some incipient failure warning characteristics. Both grease- and oil-lubricated bearings have been used in rotor hinges. The um of oil is favored for the majority of cunent helicopters. Some characteristics of the two systems are: 1. Oil Lubrication: a. Oil is satisfactory over a broad temperature range and is changed easily in ra*.onse to environmental changes. b. Oil sight gages provide positive indication of lubricant presence. c. Oil permits "on condition" maintenance. d. Reservoirs should be located so that centrifugal force drives oil into the bearings.
K A
r
{-
JV S-AMCP
706-202
c. Reservoirs and lubrication cavities should be refillable without the necessity for venting to avoid air pockets. f. Oil rcquires elaborate seals and the maintenance of seal integrity. Radian-positive-pressurm seals are commonly used for dynamic sealing. These seals should be installed so that contaminants do not enter them through centrifugal i'nrce. 2. Grease: a. Shields or simple seals are adeqiate when grease bearings are rclubricated at regular intervals. Grease retention is fairly gocd with a failed seal. b. Purged grease tends to exclude external contaminants from the bearing, c. Grease in oscillating bearings tends to channel, and the soap base may harden. Regreasing mL'y be ineffective bccansz the hard soap may prevent proper distribution of new grease. Premature failure may result. Yet, if channeling occu-s, debris detection will not oc an effective means of faijure detection. d. Changing greases (as for extreme low-ternperature use) can be accomplished effectively only by disassembly. ".-•-..1.1 CVlihidcal Rolleir itearlogs "
Cyiindricai roiicr .
.
.
i
&-.c
high
,i;ol
capacity withi small space requirement. Design considesations mainly are ,mpirical, and many of the factors are discussed in detail i,1 Ref. 42. One method of computing the basic load capacity with oscillatory motion is given in Ref. 43. Other factors must be determined by endurance testing and service experitnic. Some of the factors that influerce the life of cylindrical roller bearings in rotor blade retentions arc: I. Angular deflection of the hinge pin, which can cause concentrated load on one end of the rolles 2. Crowning of the roilern, which can pfovidt a better stress distribution and tend to ,ninimime the effect of hinge pin deflections 3. Roller guidance (e.g., use of a cage), wh;ch can reduce ro;ler misalignment 4. Large angles of oscillation, which can increase the number of stress cycles on the bearing rollers and races 5. Mounting fits, which must be as specified for particular application in order to obtain rated capacity 6. Type of lubricant,
Sof
5-5.2.1.2 Tapered Roller Bearings Tapered roller bearings have very high radial and thrust capacities. Design factors are similar to those cylindrical roller bearings and are discussed in Ref. 42. These bearings can be mounted in pairs to pro-
vide a capability to react moments as well as forces. The roller taper direction of paired bearings can be reversed to increase the tolerance to misalignment. 5-5,2.1,3 Angular Coatact Ball hearInp An angular contact ball bearing carries a high, onedirectional thrust load in combination with radial loads. The end faces of these bearings can be ground so that two or more beatings in tandem will share a thrust load. A common method of pitch (leather) hinge construction is to use stacked angular contact bearings to react the blade centrifugal force, with a reversed bearing to preload the set and to react thrust reversal. This configuration has the capability of reacting rn-omenms and radial loads as well as thrust. Design factors for these bearings are aiscussed in Ref. 44. 5-5.2.2 Teflon Fabric Beatings Teflon fabric bearings are used in the main hingcs of several operational helicopters. e.g., in the 0-6 lag hinge and AH-IG pitch bearing. These bearings are even more widely used in rotor control systems where they have demonstrated their ability to withst-nd hi;h I'cds in an adw vs environment. The Teflon fabric liner varies in thickness from approximately 0.01 to 0.02 in. deperdirrg upon the manufacturer. Strands of a material such as cotton, Dacron, or Fiberglas are interwoven in the back surface of the fabric. The fabric then is bonded to the outer hou.-ing, with the non-Teflon strands providing good bond adherence. Various bonckng agents, bonding procedures, and fabric weaves cre used; these can result in different characteristics of' the finished bearing. Design considerations for these bearings are: 1. Loads a&fdl inotion.s. The cfftl of loads and motions on Teflon bearing lives is based upon empirically determined factors. It is general design practice to compare bearing fives as a function of PV, where P - pressure, psi, based on the projected area of the bearing surface in the direction of load, and V average velocity of contacting surfaces, fpm. The acceptable value of PV for a given life varies with the pressure. Also, load reversals may reduce the acceptable value by half. Large-diameter bearings appear to withstand a higher PV.level than do small bearings. It is clear that PV is only a convenient index for comparison of bearings in similar appli.ations, and real design values will depend upon endurance
test data for full scale bearings. 2. Friction. Measured values of the coefficient of friction of Teflon-fabric bearings under loads comparable to those of rotor hinges gcnerally are in 'he 5-31
A\
reage of ju 0. 1 to 0.2. Contamination of the A bearings in service has reultod in higher values. Hig uniutwn may add significasidy to rotor system lWads and should be considered in the design. Radial bearing fo.-ces due to differential expansion also should be considered as a possible source of damaginS frictional loads, 3. Wear characteristics. Teflon fabric bearings operating at a given load level will wear at sn eswntially constant rate, up to approximately half the liner thickness. Bearings should be rep.laced at this time.
Al
The we"r rate may increase slightly when the backing
material is exposed. When loads apt not reversing, bearing wear may not always be reflected in increased clearance. Wear debris collects on th.- unloaded side of the bearing so % that the fit appears to be tight. Unless it isknown that clearance will increase with wear. means of wear other than checking clearance, or *deterniination lightness, should be planned. 5-,U.3 Flexlng Elaemens I.-.The centrifugal force on the blaide acts as a thrust Anr
nth np hinge Flexing tension elements are
used in a number of helicopters to react thia fomx while also accommodating movement In pitch. The two mtst commen types of flexing elunents are metal strap tension-torsion assemblies, shown in Fig. 5-23, and wire-wound tie bars, shown in Fig. 5-24.
thickness, and uas*="ie with and without spacerss sparating the straps.. suetnson-toruaon assemblies shown have straps of 0.032 in. nominal thickness stainless Nteel, with slots as shown to reduce the stresses due to torsion. Thin shims separate the straps at the ends to reduce fretting. The assemblies provide the capability Cot *43-deg blade motion wider design loads, and in nonrol4 operation arm cyclod approximately :L6 dog duaring each -otor revolution. Torsional stiffness of the stamp aaaembaes does not affect the pitch control forces sirnificantly. For example, the larger assembly of Fig. 5-23 has a torsional spring rate of 120 in.-b per deg under the 85,000-lb centrifug-l load of the blade. For the highspeed flight condition the increm ent of pitching moment contributed by the tension-torsion assembly is less than 3%of the total predicted blade pitching moment. The strap assemblies of Fig. 5-23 have shown excellent fail-safe characteristics. Fatigue failure is charnctc.-ized by breaking of a single element of a
strap, senerilly on an outside corner, followed by ividluas p itabsaon to otr outer eeet fe much continued cycling. (3round-air-ground cycle testing also has resulted in slow failure progression from element to element in the lug.. The tension-torsion assembly of the OH-6A heliis shown in Fig. 5-11. This assembly provides
5-5..31Tealeaterlo. tra Asembiescopter ".23. Tes~ataison tra Asemibitsflexibility for flarping as well as for pitch, or Many configurations of tension-torsion strap feathering. The IS stainless steel straps carey the conassemibliw~ have beeni u"A~ in addition to those shown trifugal loads from one blade lag hinge across to the in Fig. 5-23. These include different slot oppositt hinge and provide a fail-safe. retention arrangements, unslotted straps, straps of uifferent system.
"5-.2.3.2 Wire Th-Ar Asa mles
Q *
-.
CHAI
0
Very-high-strength. am'all-diamecter wire is wound around end fittings to form a lightweight retention system that is flexible torsionally under blade centrifugal loads. The inherent torsional flexibility can be varied over a range of approximately 10 to I if desired, by --hanging the configuration (Fig. 5-24). Normal torsional stiffnesesM like those of the flexing strap assemblies, are such as to have insignificant effects on control loads.
"5-.2A4 Elastometic Bearings CNA
0 Flwe 5-23. CH-46 and CH-47 Temlon-Torsion SthP Assemblies 5-32
Elastomeric blade-retention bearings are based on the principle that a thin layer of clastokner will withstand high normal (compteasive) forces and still permit high shear deformation (strain). By using alternate layers of elastomer and metal, the blade hinge( forces can be carried as compression of the elmstowmer and the hinpt oscillation carried as shear.
AMCP 706o202 Among the advartge, offered by clowtomeric bearings are: i. Elimination of lubrication requirements 2. Improved maintainability and reliability 3. Sand, rein, and dust resistance 4. Compressive loading, giving the ability to carry loads after severe dqradation (faii-safe) 5. Surface deterioration as the normal form of wear, giving visible failure warninS. Some of the morm common configurations of elutomeric bearings are shown in Fig. 5-25. The cylindrical bearings for radial load and the thrust bearings have bemn used to replace conventional rol!ing elcment bcarings as blade retention components. The spherical elastomer permits complete blade articulation - pitch, flap, and lag - in a single bearing, while reacting the blade centrifugal force.
(A) RADIAL BEARING
ALTERNATE
ELASTOME R IC AND METAL LAYERS
-~
~N1~(B)
THRUST BEARING
(A)FLEXIBLE TIE BAR ASSEMBLY
ALTERNATE SPHERICAL ELASTOMERIC ETAL LAYERS
(C)SPHERICAL BEARING (B)STIFF TIE BAR ASSEMBLY
)y
T
Flgure 5-24. Torsionally 'Stiff'sad 'Fiexible Wirt-wound Tie-ar Assemblies
Figure 5-25. Elastonewic Ikarings
3
!L
*
I/r
.
AMCP 70O-2O2 Stops 5-5.25 La Dampeirs, Laend-a The lag damper of an articulated rotor must meet blade stability requirement& in ground resonance (par. 5-4.3) and in flight (par. 5-3.6). Two common means of energy absorption in lag dampers are hydraulic shock absorbers (used on the majority of large hcli&. pters (Fig. 5-26)) and friction dampers (spring-load oscillating disks, used in several small helicopters). Although timpler, lighter, and less expensive, friction dampers gSnerally arc less reliable. Therefore, the use of friction dampers in new rotorai is discouraged.
TUflon fabric bearings frequently arc used for mounting lag dampers. For satisfactory life with reversing loads, PV values (par. 5-5.2.2) should be approximately 1/3 those found satisfactory under nonreversing loads. The CH-53 has auxiliar, pistons in its hydraulic lag dampers. The pistons ar,. pressarized to force all blades to the lead stops so that vibration during rotor startup as minimized. Hydraulic lag dampers frequently are used as lead and lag stops. Integral hydraulic cushions can be used to reduce the impact of the blade agtinst the stop. Principal design conditions of the lead and lag stops (whether integral with the damper or not) are: I. Predictable flight conditions or maneuvers will not cause contact of the lead or lag stops. 2. Lead stops will not yield due to rotor brake application. 3. Lag stops will not yield due to engine starting
torque. 4. Lead and lag stops should fail before any dynamic component critical to safe flight yields.
S..restoring ...... ....
I I
5-5.2.6 Droop and Flap Stops and Restrainers Articulated rotors require stops to limit the extremes of flapping motion of individual biaaca. t• Teetering and gimbal-mounted hubs have a similar rcquirement, but the motion is that of the complete rotor assembly. The droop stop must be positioned to allow normal cyclic blade motion in all predictable flight cond~tions or maneuvers without making contact with the stop. This stop position also must be 1-"gh enough to prevent blade/fuslage contact in high winds during rotor shutdown and when parked, i.e., when blade centrifugal force does not provide a radial force.
Centrifugally operated droop restrainers often are
used to increase blade/fuselage clearances. These restraincrs engage at a rotor speed approximately 4060% of normal during rotor shutdown and restrict blade droop to provide clearances not possiblc with permanent droop stops. Flap stops prcvent accidental overturning of the
blade~s in vcry high winds, and are positioned to allow
Figure 5-26. 5-34
Hydraulic Lag Damper
clearance for flight cyclic flapping motions in all flisht conditions. Centrifugal flap restrainers also can be engaged at low rotor speeds during rotor shutdcwn and, in conjunction with droop rstrain•ets greatly reduce flap hinge motion. Flap restrainers are essential for blade folding to prevent the blades from "elbowing". Elbowing occurs when a blade is folded so that its CG is inboard of the flap hinge axis; if the hinge is
'
"
I
¼
:I
t,
..... .rIi
I
7-0
AMCP 106-aV02 free to flap, the blade tip will droop unoer the blade weight. Flap restrainers can be ground support equipment if blade folding is accomplished only occasionally. Two common types of centrifugal droop and flap reattainer mechanisms are ovcenter linka3es and interposer blocks. Both of thes use weights to release the mechanism as the rotor speed increases, and springs to re-engage during shutdown, Another droop restrainer mechanism is a floating (gimbaled) ring below the hub, with projections on the flapping portion of each blade. Cyclic motion of individual blades displaces the ring to permit flapping
(feathering) acceleration. if balance weights arc used to reduce or nullify the centrifugal centering moment, the increased polar moment of inertia will result in higher loads in the control system to obtain a given change of blade .)itch. For those rotors in which cyclic pitch is obtained t y oscillation of the blade about the pitch axis, e.g., fully articulated and hingeless systems, the dynamic characteristics of the system may be important. The selection of a blade retention system should include the investigation of the response of the blade to the oscillating control force. Systems such as tension-torsion straps or wire tic-bar assemblies that have
without restraint, while the ring supports all the
known torsional spring characteristics may be
blades against collective droo,.
required to obtain an acceptable relationship between the natural frequency of the feathering motion and the rotational apeed. Control system design considerations are discussed in detail in Chapter 6. Further discussions of both nullification of the centrifugal centering moment and optimization of the natural frequency of the feathering motion are provided by Ref. 45. Arfi A Who A %
"5.3 CONTROL SYSTEM CONSIDERATIONS
9 \
In the selection and design of the blade retention system, consideration must be given to tk.e characteristics of those elements which aff•ct the loads on the rotor control system. Displacement of the blade about the pitch axis is opposed by a friction torque of essentially constant value if angular contact ball
nearinns are usewi to react tkut centrriuaai Aorc of the blade. On the other hand, both elastomeric bearings anJ flexing elements (tensit n-torsion straps and wire tie-bar assemblies) have the characteristics of torsional springs. When the elements of this type arc used to react the blade centrifugal forc, the torque opposing angular motion of the blade is proportional to the displacement.
For shipboard operation, or for compact stowage, it is desirable to bring the rotor blades within the fusolage envelope. When the blades are stowecd in this manner, there is much less possibility of blade damage when moving the helicopter, or when other helicopters or vehicles are moved in the vicinity. Blade folding is preferable to removing blades; there
In addition to torque resisting motion about the
is less chance of handling damage, and rctracking can
"pitchaxis that originate in the blade retention system,
-
r',3.35
there also are torques that dspend upon the inertia characteristics of the blades. The first of these, known variously as the propeller moment and the centrifugal feathering moment, is a torque that opposes displacement of the chordwise principal axis of the cross section of the blade out of the plane of rotation. This torque is directly proportional to the difference between the moments of inertia with respect to the principal axes of the cross section and also varies directly with ill. The presence of the torque also has been referred to as "the tennis racket tffoct". For tail rotors and for main rotors of small helicopters for which the rotational speed is high, it may be desirable to nullify this torque by equalizing the moments of inertia. This may be accomplished by the addition of appropriate balance weights at the root of the blade. The second torque that depends on the characteristics of the blade is the conventional inertia reaction, proportional to the polar moment of inertia with respect to the pitch axis. This inertial torque varies "with, but is opposite in direction to, the pitching
LD
l r@Laf9lfN FY.4
be avoided. Either manual or powered blade folding systems can be used, depending upon operational requirements. Consideration should be given to blade folding in the initial design of a rotor system, even if the basic heilcopter criteria do not include this as a requiremeat. Appropriate decisions as to the method of blade attachment, radial location, and clevis clearances of the blade attachment joint will simplily incorporation of folding at a later date. 5-5.4.1 Design Rtquireaseiut 5-5.4.1.1 Manual Blade Foldlag The following prvisions are required: I. Rotor brakes, rotor locks, or means of securing trc blades to the fuselage to retain hub azimuth position 2. Pitch locks, locking pins, or fixtures to restrict blade motion about the pitch axis, preventing load feedback into the pitch control system forn a folded blade 3. Flap rcstrainers (articulated, gimbal-mounted,
i06~
and tKt~ring hubs). Centrifugally-operated droop ad flap restrainers are deired; if not self-contained, special ground support equipment is needed to restrain blade droop. 4. lade fold hinge, with quick, simpli means of locking and unlocking blade motion ab,)ut thic hinge 5 Clamps to restict lag hinge motion (articulated rotors) 6. Quick-disconnect fittings for attaching handling lines to blade tips (optional) 7. Acocw to the rotor heLd (steps, handholds. toeholds, or rungs) and w work platform for performing rotor-head folding operations. Major stops in a typical manual blade-folding operation of a single-rotor helicopter are: 1. The rotor should be rotated to the proper azimuth position for folding, 2. Rotor brake or rotor lock should be applied, or one blade should be secured to the fuselage to retain rotor hub azimuth position. 3. With an articulated hub it may be necessary or desirable to move some or all blades to a predetermined position about the lag hinge and to lock out any further motion about this hinge. , d . .4 4. C^^iu; . positioned to the proper setting for folding. 5. Pitch locks should be installed on all blades to be folded. 6. Flap restrainers: a. For articulated hubs without automatic flap restrainers, a flap restrainer should be installed on all blades folded 90 deg or more. b. For teetering or gimbal-mounted hubs, flap hinge motion should be locked out. 7. Racks should be installed to secure the folded
structure, if required. blades to fust.lage or other .
a. A blaud-suppuning polc and steadying f inus as
The following componGnt requirements are necessary for a power blade-folding system: I. Power rotor orientation mnchanism 2. Rotor lock to maintain azimuth position 3. Automatic droop and flap restrainers (articulatcd, teetering, or gimbal-mounted hub) 4. Contro! position indicating devices for pilot 5. Power pitch locking device 6. BI. le lag hinge positioning device (articulated hubs) 7. Blade fold hinge unlocking devicc 8. Blade fold actu~ators.
The power blade-folding mechanism for a typical blade of the CH-46 is shown in Fig. 5-27. An electromechanical actuator housed within each rotor blade folding hinge pin operates a linkage that sequentially inserts pitch lock pins, positions the blade about the lag hinge, releases the blade fold hinge lock, and rotates the blade to the proper fold position. 5-5.4.2 Operational Requireantats Blnde folding frequently must be accomplished in an adverse environment with poor lighting, winds, rain, and possible helicopter motion. hladP fnldina under these conditinns can mnial M result readily in crew injury and human error. In view of this, demign of folding components should consider: I. Minimizing loose components 2. Minimizing large or special tools 3. Attaching flags to all fixtures (pitch lock pins, etc.) so that it is apparent the helicopter is ursafe for flight
4. Providing adequate access and working areas on the helicopter for performing the folding operations. *•l~orut•.• b~twu• n r, A-I• Notice andi betww•n
required by blade size and accessibility, should be atand tached. The blade-fold joint should be unlocked, each blade fo!ded and secured. Blades may be secured in racks, by lines to helicopter structure, or about the fold hinge, 9. If the tail assembly is to be folded, it may be necessary fold itinbefore the cycle.5-. main blades or at an intermfediate topoint the fold
blades and helicopter structure should be adequate to allow for blade, hub, and drive system deflections under wind loads or due to motions of the helicopter. If the blade-to-fubelage or blade-to-blade clearances are inadequate provisions for biade i,..-ks or bladt securing lines should be matc.
"-o4.1.2 PoWi
Blde Foldif For flight safety it is mandato'ry that a power blade folding system be properly interlocked to prevent any malfunction from occurring in flight. Interlocks also are necessary to prevent damage to helicopter comnportents due to improper sequencing of the power blade-folding systen. it also must be apparcnt to the flight crew that the blade unfolding sequence is corn-
cussed in relation to design and operational requirements, the entire blade-folding operation should be reviewed from the system safety viewpoint. The identification of potentially hazardous conditions should be made from the viewpoint ofa material failure/malfunction, environmental conditions, personnel error, supervisory influence, or any combination of these factors. Maximum effort
plete and the aircraft is safe for flight,
should be made during the design phase to reduce the
5-36
I
S e SftCe erte In addition to the specific safety consideration dis-
-
AMCP 705-202 hazard of these failure modes. Guides for this safety
must incorporate strength and stiffness character-
analysis include MIL-STD-882 and Ref. 46.
istics that will meet the applicablc structural design
.- 6 ROTOR BLADES 5-6.1 GENERAL
characteristics, and will do so efficiently and econumically.
As described in par. 5-2. design of a helicopter rotor involves the dctermination of optimum values for each of a number of parameters, i:cluding those that define the blade gcometry. The blade designer
The blade geometry is defined by the parametcr of twist, planform, aud airfoil section. Selection of values for these parameters generally is accomplished during preliminary design. The type of parametric analysis required to optimize the rotor design
criteria and also will provide acceotable acroelastic
FOLD HINGE
-
PITCH HORN
2
b)!tW
aI.PITCH
-7-1j
HORN LAG P11CH
FLAP HINGE
LAG H'NGE'-
HUB
POSITIONER
0
•,:•
•.,PIN Figire 5-27.
"BLADE CLAMP
FOLD HINGE
BLADE FOLD MOTOR INFOLD HINGE
CH-46 Power Blade Folding Mechanism 5-37
AMCP 706-202 is described in pEr. 3-4.1, AMCP 706-201. The considerations pertinent to the three principal flight conditions - hover, high-speed lift, and maneuvering are reviewed in par. 5-2. Ir the paragraphs that follow, the sign-ficance of the two types of design parameters - aerodynamic and structural - is reviewed, with emphasis on the detail design and :manufactuie of rotor blades that will satisfy these requirements. 5-6..1
Twist
SGenera'y, rotor blades have a linear twist on the
Sg.ad .-
.
order of 4 to 8 deg, and the blade tip angle of attack is less than that at the root (washout). The primary considerations leading to se:ction of the deaign value of twist occur in the trade-offs between hover efficiency and the delay of high-speed, retreating blade staql. Thfor SThus, f raa particularraircraft aic at mission m si n profile p oietthata combines both hover and cruise, an optimum twist must be determined that will allow both high hover i efficiency. cThe ~~~~gross weight and goodd cruise The effect of twist in the hover Dode. is to .reate a more uniform intlow distribution trom the biae tip to the root. The so-called "ideal twist" (which results in unretical w lus of twist n-ar the blad, rnfo ) theorctically would result in a uniform inflow distribution across the rotor. Large an! ounts of twist, up to 12 deg, approximate this distribution over at least the outboard half of the blade. Twist has the effect of reducing both the induced gad profile drag losses of ,.he rotor so that the hover efficiency, generally refer:-ed to as Figure of Merit, is increased (-cc par. 32, AMCP 706-201). The theoretical maximum value for Figure of "Meritvadue is unity. This value can occur only if the rotor has no tip losses and also posse2.ses no profile drag. These conditions cannot occur, so an actual rotor Figure of Merit value always will be less than unity. The effect of twist in the forward flight mode of rotor blades is to lower the pitch at the tip while maintaining a larger angle near the root. This reduction of tip angle of attack gives a corresponding decrease in the rotor profile drag power, which, in turn, allovs a higher forward speed to be obtained, parametric studies used to optimize design twist for a particular aircraft also must include torsional deflection in obtaining the section angle-ofattack and the corresponding aerodynamic loading, Effects of drag loads and centiifugal twisting also should be included ;n the elastic twist angle deterruination. Th: forward flight angle-of-attA;k determinartion also should include the effect of blade pitch rrte (tennis racket effect) on the instantaneous twist
S.Aerodynamic
angle as it varits around the azimuth. Commonly the inertia contribution -- i.,., ccr.trifugal twisting due to blade pitch and pitch rate - will be greatar than thc aerodyninmic twist.ing monrents, and the net tarsional deflectiou is in the nosedown direction. A further requirement is that at high forward-flight speeds, the advancing 'blade tip has zero, or near zero, lift load in order that the corrcsponding high Mach numbc- drag be minimized. For a given amount of twist and a given forward speed, this minimum drag can he met orly at one specific gross weight. For operations at gross weights above this value the increase in required collective pitch resu'ts in increased blade-tip lift and drag loadings Also, at gross weights below this minimum, a highly twisted blade tIp operates at high forward speed with negative lift on part of the advancing side of the disk. This causes aen nosedown eo e mcontrol-load yb i i - pulse, b c ueandhther aircraft ul ntvflight r envclope may be limit-d brcause the resultant vibration exceeds prescribed limits. optimum amount of twist that isloi etitdbfor forward h lgtpwrrqieet flight power requirements also is restricted by the relatively linear increase in oscillatory flanwise hending moment with increased twist. In general, the powe," consumption and blade torsional moments duc to compressibility effects can be minimized if, at the design condition, the blade is twisted to produce zero lift on the advancing tip. 5-6 1.2 Planform Taper As with twist, the effect of planform taper is to give a more uniform inflow distribution across the'redisk during hover and thus to increase the Fs& of Merit. The local induced velocity is proportional to the square root of the blaied ie-_e.inn lift, which in turn, is directly proportional to the local blade chord. Thus, by increasing the root chord over that at the tip, the induced velocity over the inboard portion of the disk can be increased, simultaneously increasing the thrust over the inboard portion of the disk. Experiniental results have shown, however, that the oscillatuy bending nroments are increased as the planform taper is .ncreased (i.e., tip chord is much less than root chord). The higher cost of producing planform-tapered blades has ruled out their general use. In addition, a blade with planform taper requires a thickness taper in order to retain a uniform airfoil section with known characteristics. Also, significant planform taper results in a srmaller blade tip cross-sectional area available for tip balance weight placement. Also, excessive amounts of root chord - as dictated from Figure of Mer;t optimization studies - can cause a premature power limit on forward speed due to an increase in profile power.
.
S~AMCP
5-6.1.3 Airfoil
w
msSe..lon
In addition to the usual need for high lift-to-drag ratios, stall angles, and critical Mach numbers, rotor
blade airfoils require low pitching moments. Airfoil
cfficicnts. However, this bcneficial effect disappears
pitching moment coefficients that vary appreciably with angle of attack give periodic pitch link inpats that are undesirable and that, in turn, can lead to periodic forces and vibrations. Thus, although in forward flight the angle of attack varies with azi-
as Mach number is increased. A further benefit of the delay of compressibility effects due to the use of thin, cambered airfoils is that noise levels due to these effects are, in general, decreased for a given flight condition. The noise
muth, it is desirable that the blade pitching moment coefficient not vary. Usually, it is preferred that the pitching moment coefficient be zero so the corresponding loads do not vary with the variati3ns of local airspeed.
generated by a blade intersecting a tip vortex is affected only to the extent that the vsrength of the vortex is affected. This particular form of noise is caused by a .railed tip vortex being intersected by the following blade, with a rcsukhnt rapid change of
The usual starting point in airfoil selection is the
• \
J
706-202
may have undesirable effects on blade profile power. At the same time, cambered airfoils extend the low Mach number, retreating blade drag divergence boundary to regions correspording to greater lift co-
minimization of rotor power requirements for th design cruise and hover conditions. Two-dienensional airfoil drag data at the design lift coefficient arc used in this determination. The static variation of drag coefficient with Mach number, along with the change in the drag divergence boundary, also i- used. The drag reduction potential of thin airfoils is well known, and has the greatest effect near the blade tip
-.
.,,,,,
angle of attack. The corresponding pressure change
causes the slapping noise that is characteristic of heli. copters. At a given flight speed, a lower lift-to-drag (L/D) ratio will reduce the vortex strength and hence lower the noise. Several airfoil sections that ,:re used or could be used in helicopter blades are shown in Fig. 5-28. The main geometric properties of these air foils - such as thickness ratio, leading edge radius, and camber -
due to thc higher Mach numbcr environment therc.
are identified in this ligure.
For ease and economay of manufacturing, a thin air"foilat he blade tip usually is auhieved with a uniform
It, addition to the characteristics shown, some blades possess a thin, teailing edge extension strip
root-to-tip thickness taper.
that extends beyond tWe "true," airfoil trailing edge.
On the other hand, the ret. -ating blade stall and drag divergence characteristics of thick sections are superior to those of thin sections in several series of airfoils. In the low Mach number region of the disk, the thick sections allow a higher lift coefficient to be cbtained before the onset of drag divergence. However, the advantage of these 4"! ck sections is reversed in the high Mach nuinbec environment. Tnus, the airfoil section characteristics for the advancing and
In the usual blade, this strip 'sosed as a base to whi..h th. upper and lowrr skins are bonded. The strip can be tailored in length and thickness, or number uf laminates, to obtain the dcsir'd edgewise stiffness and fatigue properties. Airfoil characteristics shown in Fig. 5-28 also can be used joirtly so that the d-si, ,nle properties of seve.el characteristics can be incorporated into one airioil. For example, a blade design could be based on a thin airfoil with e dtoop
e
retreating blades are in conflict,
nose. This would combine the benefits of reduced
The addition of airfoil camber ana increased leading edge raduis tends to improve the low-specdstall characteristics of the symmetrical airfoil sections commonly '-sWd .or rotor blades. At high MAch
drag an the hi3 rhMach number advancing blade tip iýi:h increased maximum lift coeflicient on the retreating biade. This hypothetical blade could be modified further with a iarpe leading edge radius, and
numbers, the eftect of camtrcr is ;o decrease the
could have its aft section produced with straight
maA'muni obtawable lPit coefficient. However, this effect is riot too significant because low lift coerficir --ts are desired in the advancing blade, high Mach number region. Camber generally is applied to the forward portion of the rotor blade airfoil cross section in order to retain low pitching moment coeffi cients. This leads to the "droop snoot" terminology used by at least one contractor. In some instances, the trailing edge is reflexed slightly to counteract an otherwise unavoidable amount of pitching momenz and the corresponding cyclic control loads. This method of eliminating undesirable pitching moments
"slzb" sides. These changes would improve, respectivcly, the abruptness of the blade stall chatacteristics and the ease of manufacturing khe blade aft section honeycomb er web structur-e. Further, the machining of the main bonding molds for a slabi ded blade will be easier, hence, less expensive. The decreasz in blade flapwise, edgewise, and torsional stiffnesses ca.sed by this particular geometric shape could be restored with the proper slection and layups of advanced materials such as boron or graphite composite., but the slab-sid,:d airfoil may r.t provide as high a value of LID. By ,use of this type of 5-39
&.............
trade-off procedure it should be possible to obtain a blade airfoil that represents an optimum configuretion for the specific hWticopter mission. Blade !ip geometry has been found to ha~vc impor-. tant effects upon overall rotor performance. Early studies genr-ally used constant chord sip covers wiht flat, stmiround, or other types of curvature. These studies indicated thiat limited perfomiance benefit was obtainod with tip covers ftht had xomplicated curvature and were thus difficult to manufacture. The major performance gain of these tip covers. as indicated by the rotor lift to drag LID uhually could be ti accd to an increase in rotor radius. More recent studies have cicplered owept leading and trailing edgecs in aim, effort to reduce the tip vortex velocity and strcnjihI. These results have shown that a planforni
THICKNES PATIO. /
M&XiL*A~ IHICKNES$ I
K
LOCAýAN OF
C. HCNSS
ILEAOIIG FADIUS
-
ufAN CAMER'.JNE
_-I--
-
.---
I~
____ r~Y
2
J
-
-
c~c;;7 ~Z .
______on
_______IO
U"UIN
MA~IL4IAMUc4
CAUSERfoil --
STANARDAIRFIL-ACA 012a -
~
THIN AIRFOIL-MACA 0006 Sis
S-~ SHARP NOSE AIRFOIL-NACA 63AI,,
~
-
.chango,
Q_~z~ SL-AB SIDE AIRFOIL -*---..on Sin
CAMBEIED AIRFOIL __________________________edge S~coupled
DROOP NOSE AIWOIL
Flpure 3-28. 5.40
aneuivers, the local conditions
~in the following
bade tip are changing rapidly due to &hevortex effect the angle of atiack. Therefore, this radial segment of blade dies not behave as two-dimensional wind tunnel data wouald suggest, and time-dependent airsection characteristics should be used itn the anaof acerodynamic loads and roto'r perforniance.
*---------lysis -~
-
taperwd bld,,tip does tad., Am hitp vor%%Iand henoe both tlim corm~ponding naise and tho UWAI~laory loads origiatisng rsear the tip of the blade. Inadilon. the aft sweep of the leading ed*r delays the effoits of comnpressibility, as With v fixed wing. This allcowa an increase in forw'ard spood fo.- fixed values of rotor spc*- and available power. Other attempts at bLkde tip gvometry mnvdification to deecrase the tip vortex, such av~ installing slots ur holes, have not proved successful. The tipplicability el stundord,. two-dimensional airfoil data in rotor ana~ysis is q~tcitionablc whenever the shed tip vorte-x al* progiahes the following blade. EBecaus-. the bled%: tip gcomctry has such1 a stong influence on the tip vortex, the allawfible spatial relationmhip betwoeen the vortex find blade for future high-speed heqlicopters should Ix analyzed for various tip configurations. Recent flight tests have indicat~d that thc oscillatory airloads tend to be concentrated at the bladc tips. Most harmonics are characterized by higher lobding at the tip due to the ,ipulsivc ntature of tip vortex interference. In tiousteady
TypkMa Helkcoper BoIrM ade Airfoils
Two-dimensional static tests show that r4te low Mach number stall of thick airfoils of cvrtain series is leading-edgec phenomenon. This stall is characv'rized by a sudden sepafation of flow over tht entire upper airfoil surface, and resufts in ani abrupt change in the lift curve slope. Also associate-S with this effeci an instantaneous nosed&.wn pi-iching mom!.rnrt. No warning occurs that would ini-'aite stall a i.-riminent. This instantaneiu!, OTfet docs. not occur ii- two-dimecnsional tests of thin airtoils. Rather, a gradual in lift curve slopt; mkatý r~oce, alcog w~th a more gradual increase int nosedown pitching .noment due to stall. Many investigaiievs have been cunducted as to the c~ffects of noststeady acrodynt.PU*CS various airfoil sections. These tests indicate that, the dynamic enivironirint of irtcret.'ni angler of tvLtack with time, stall as indicated by a iom K- Ui~t Ui not occur in the manner pnedictcd fiem- static tests. When a sharp loss of !ift does ozcum due to Iai4 separation, the resulting impcac
is highly
to the dynamic response of fl~t blade in toision. For instance, a fixed itupuls: due to L 3uchlw loss of lift on the :etreating blade niuy hrtv.- r.Sitter effect on a rotor sy! 'em haviing both at b4w cowmrol stiffness and a torsionally soft blade. Tot torn it!
1.
)AMCP7W * *
*•
stiffness of an airfoil cross ection is approedmately proportional to the square of its area for a Sivrn type "ofcell constru-taon. Thus, a thick airfoil can be manufactured with higher torsional stiffness. Therefore, for the airfoil series with which the thick section receives a sharp impulse due to sudden stall, the torsional response of the blade may be leks for a given amount of cointrol jystan stiffn%4 than when a thin section is used. Dynamic pitch te.its of various airfoils have shown clearly that the increase in the maximum lift coefficient is large for low Mach numbers and decreases as Mach number is increased. Thus, as the retreating blade increases pitch, it can ivacn a greater angle of attack before stall occurs than that predicted by twodimensional airfoil tests. The amount of this increase is dependent upon the airfo'l geometry, particularly the amount and location of camber and the leading edge radius. The lift coefficient that corresponds to the drag divergence angle of attack obtained from two-dimeasional tests, however, remains an adequate indicator for oscillating drag divergence. The more gradual increase in pitching moment due to the
bladc tip vortex strength and trajectory in conjunction with the objectives of obtaining a high hover Figure of Merit and a low value of cruise power reiuired. Blade twist, chord, and thickness, and the corresponding physical properties of the blade, should be chosen to minimize the responses due to vortex action and aerodynamic hysteresis effects, in addition to resonant conditions. These types of anslyses should include the effects on the blade of the entire control system, as well as possible shrft, or pylon bending. Dynamic blade stall effects at the first torsional natural frequency and at the once-per-rev rotational frequency of the blade should be included in the blade response analysis. 54.2 BLADE CONSTRUCTrION Rotor blade structures may be broken down into three major elements: the spar, the aft section (sometimes referred to as the fairing), and the root end retention. Secondary elements are tip closures and hardware. trim tabs. and tuning weights. While wooden rotor blades still are in use and probably will be used to a limited extent for many yeaus, they are
couid result in Itrgc losses of lift if these torsionally softer blades untwist sufficiently to precipitate a loss o! lift. Sharp increases in torsional loads then occur and can result in the same net effect as leading-edge separation of thick airfoils. It has been foomd from several experimental sources that cambered airfoils with a slightly increased leading-edge radius possess superior pitching moment delaying characteristics. A large leading .edge radius also helps keep the blade section CG iorward, which delays the onset of both pitching moment stall and classicdl bending-torsion and stall flutter. i-or airfoils with sharp leading edges and with their maximum thickness further aft, the
tcr development and, therefore, are not addressed in this handbook.
overall blade CG can bt moved forward with a large
This extrusion may be a "D" spar, usually con-
tip-over balance weight. This is not as desirable as obtaining a more uniform forward CG position from root to tip, since individual 31adc radial segments may still be acted upon by undesirable moments.
sisting of a single cell, or it may be of an essentially trapezoidal shape. The "D" shape conforms to the forward portion of the airfoil shape, with the vertical part of the "D" serving as & shear member. The
A form of negative damnping can occur when the blade twist rate and the loading due to the acrodynAnic pitching moment are in the same direction. This may !kiadto excessive torsional response and to subsequent loss of lift on the retreating Htade, chnracterized by excesive flapwise bending amplitudes. As with torsional stiffnew, a thicker blade obviously possepss more flopwise stiffness, for a given type of
hollow trapezoida spar, sometimes referred to as a box beam, may be made to conform to the upper and lower sides of the airfoil surface but requires the addition ot a shaped component on the forward side to provide the nose rdius of the airfoil. This shaped componeiit usually tervol also to provide chordwise balancing of ttc blade and therefore is made of.brass or some other relatively dense material,
€onitruction and hance will respond less to a fixed amount of flapwise lift input. Determination of blae geometry shouid consider the effocts of nonoteady aerodynamics along with
54.2.1.2 Solid Extruilog This extrusion iR solid in the sens4 that its crosssectional outline may be tiared without lifting the
56.2.1 Spar The major load-carrying member of any rotor blade is the main qpr, whether it be designed for structure only or also as a part of the aerodynamic shape of the blade. It may he of monolithic construction or may be assembled from two or more components. The predominant types of spars are described in subsequent paragraphs. 54.2.1.1
Hollow Ex•taloa
=
.:
".'2.•
.~ •
-- :+
"%::"-
. ::•
AMCP 76-202
"tracingstylus. It may be referred to as a "C" section,
__•ters,
opening tcward the trailing edge. As with the 'D" section, it conforms to the forward portion of the airfoil shape. Commonly the wall is thickened a considerable amount at the nose to provide chordwiue balance and resistance to impact damage. The "C" section may or ma) not be closed at its ai. end with a separately extruded or formed shear web. The principal advantage of extruded aluminum spars is relatively low cost in production. The "ID" and box beam configurations lend themseh'es to internal pressurization as an in-service inspection system for cracks. The advantage of the "C" section is that its internal, surface may be inspected during manufacture. A major disadvantage is that the use of extruded spars is confined to constant-section blades. An added disadvantage is the poor resistance of aluminum to erosion. In low-performance heiicopwhere changes in aik'foil nose radius are not critical, this problem car, be ignored; however, it usually
Reliability is enhanced in that the quality of the raw material can be closely controlled and inspected prior to spar fabrication. The raw material does not undergo any fundamental change during the fabrication process. 5-S.2.1.4 Round Steel Tube One of the earliest types o" spars for rotor blades was a round steel tube, and certain advantages still exist. Obviously, a touid tube cannot be used to constitute a part of the airfoil shape, but must be buried withia an enclosing structure or envelope. Inherent in the various process= for producing such spars is the ability to taper both diameter and wall thickness continuously or in smooth steps - providing considerable latitude in stiffness, mass, and -aerodynamic taper of the rotor blade. The heaviest portion of the tube protrudes from the root end of the blade enivelope, and may have integral attachment lugs or may simply be a cylinder that accepts a socket type of
is necessary to cover the aluminum with an erosionresistant shield at the leading edge of the blade. From the standpoint of efficient design, the fatigue
retention fitting. Generally, excellent material properties are obtained in tubular spars due to the nature of the cold-
tive. 5-6.2.1.3 Formed Sheet Metal This type of spar is fabricated from mu•.iple components, the minimum being a "C" section and a shear web. The shear web may be the web of a channel section, with the flanges providing surface area with which to bond or braze the channel into the "C" section to form a "'D" shape. /.dditional webs may be added to made a multicell structure. In most instances, a continuous or segmented balance weight
this proce.ss is accompiished with sufficicni pi io,-, to avoid stress raisers, high fatigue strength can be . obtained. Further, with proper blade design, the material surrounding the spar tube can have sufficient independent strength to make the structure highly redundant. A disadvantage of the completely enclused spar is the difficuity of access for inspection, 5-6.2.1.5 Formed Metal Tube An alternative to the round tube is the formed metal tube. Generally, this starts with a round tube
i*i
w'hich
strenge!h-to-weiaht ratio for aluminum is not attrac-
;In or
a the nose radliiuc ,f the. 'C'` and
may contribute to the overall structure, particularly for chordwisc stillnesm. There is a wise choice of materials for fo-med sheet-metl spars, ranging from low niloy steels to any of several types of stainlesrs steels tyr noaferrous al!oys such as beryllium coppr. Among the advantages of this type of coistrucilon is the ability to taper the spar in almo:t any manner desired. Another is the ability to tailor the gages of the different components to achieve a given set of stiffness and strength requirements with greater precision, Perhaps the greatest advantage is the redundancy of the structure. The bondlines between the components are effective crackstopper6 so that, even if the "C" spar should fail, the remaining structure can be designed to carry the loads and prevent a catasti ophic failire of tht blade. Finally, depending upon the alloy and tlh-configuration, thc spar can provide adequ•te erosion protection without n extra shield, 5-42
working process employed iii their fabrication. When
-- bý,ýtu-ntlv
6
fnrm to
either a "D" or an
oval shape within the blade envelope. In the former case, the "D" is the forward portion of the airfoil contour. In the latter case, the oval tube is encased within the envelope of the airfoil, much as with the round tube. The oval shape permits a thin airfoil compared to the original tube diameter. In either case, taper of the airfoil is quite difficult to achieve, athough the wall thickness can taper so as to give the desired mass aad stiffness distribution. The root retent:on alternati.c4 Ft, identical to those for the round tube. Most of the advantages and disadvantaaes are the same as for iound tubes. 54-2.1.6 Molded Reinforced Plastic Molded reinforced plastic lenrls itself to alii,st any gcometric. spar configuration. High-strength fibers which may be of various types of glass, graphite, or boron - are imbedded in a matrix, usually of epuxy.
M Orientation of the fibers along the length of the spar gives a composite construction that is very strong in axial tension and is light in w'eight. One successful configuration is much like the solid aluminum extrusion. Others may be "I" beams or variants thereof. The large number oi configurations possible include a multicell section with complete shear weba molded integrally inside an airfoil-shaped shell, One of the greatest attractions of molded plastic is the ability to achieve any desired degree of taper and virtually any desired shape. Another is the ability to wrap each fiber, or filament, around t'.e principal attachment member at the root end so that there are no discontinuities in the load-carrying material. Still another is the relatively wide selection of stiffness/ strength/weight ratios that arc available through the choice of fibi-r-reinforcing material and the orientation of the fibers. A disadvantage of molded reinforced plastic is tht difficulty of repeajing with precision the properties (density, strength, stiffness) from one unit to another. This problem is being overcome, through improveimnts in the molding process. 5.4U.2 Aft Sscfie The aft section, or fairing. of a rotor blade is the aft 70-80% of the airfoil. It consists of upper and lower skins, some type of contour-stabilizing internal member (usually a structural trailing edge strip), and a means of attachment to the spar. This section may make a significant contribution to the beam stiffnesG and strengah of the blade, or, in some cases, it may serve only as a fairing and to transmit the airloads to the spar. There are many different typ's and variations. The rtist common are described in succeeding
SP~agtisphs,
5-6.2.1 Coatimous Skim Continuous skins of sheet metal or fiber-reinforces plastic may extend from the root of the blade to the tip. Regardless of the internal members, continuous skins normally carry a significant amount of the centrifugal loading and a lar, -share of thc chordwise bending anc, torsional stiffness. These contributions can be controlled closely in the case of plastic skins by the sclection of the fiber orientation. In this way, a blade can be designed to be torsionally soft and yet veiy stiff in the flapwise or cbordwise direction, or vice versa. SThe internal members that tie the upper and lower skins together and maintain the blade contour may ""'_/ be metallic or nonmetallic "'I"beams or channels, honeycomb core, foam core, or a series of individual ribs. Blades with a chord of less than 8.0 in. or blades
706-202
with unusually heavy skins may not require any internal members in the aft section. Very large blades may be constructed with individual sandwich skins both top and bottom, in which case no further reinforcement or stabilization may be necessary. Spanwise "I" beams or channels in the aft section generally contribute significantly to blade chordwise and torsional stiffness and thus are found more often in the blaces of semirigid rotor systems. Channels are adaptable as spanwise members in tapered blades since thay can be stretch-formed to the required shape. If they are bruke- or roll-formed in a constant shape, they can be placed in a skewed position within the aft section so that they follow a spanwise line of constant blade thickness. However, the use of such internal members often has the disadvantage of Complicated internal too'ing required for proper positioning, and to supply adequate pressure during adhesive bonding of the assembly. Honeycomb core as a filler between the top and bottom skins of the aft section is extremely effective in maintaining a stable airfoil contour. Although aluminum alloy honeycomb core is the most common. there is a growing tendency toward the use of nonmetallic honc)cuwb. Tht latter has the advantages of being less susceptible to corrosion, relatively resistant to impact, and - where nonmetallic skins also are employed - less suuxptible to lightning strikes. Whenever honeycomb is used, careful attention must be given to sealing a blade completely against tiie entry of moisture, because any water that enters the blade has a tendency to migrate and become entrapped, leading to corrosion and blade unbalance. Fosm core also has been used successfully in blade oft
=ttc•tni
Thp
liahtweight fnramg
reguired in this
application are somewhat more susceptible to lelamination between skin and core than arc the honeycombs, and to failures occurring within the foam itself. Generally, foam cores are pre-cured before -blade assembly. Foaming in place is !a be discouraged since it is difficult to obtain uniform quality and density. Individual ribs commonly were used with wooden rotor blades, but seldom are employed with metal or reinforced plastic blades with continuous aft sections. In the latter case, the tooling for installation of the ribs becomes quite :omplex, and contour stability is difficuilt to maintain within the weight and balance limitations. 54.2.2.2
Segnumted Skims
Blades used in fully articulated rotor systems often are constructed with segmented aft sections, 5-43
I
K4'-r, .' 4
.. . . ....
i
Straeliog
S*
o ti r•f d to as boxes, pockts, or fairinug Obviously, this type of construction provides for no centriflial load-carying ability, and makes little contribution to chordwise stiffemu unless each ofmeat is connected by a continuous, structural, edge strip. The skins of the aft section segments may be of metal or reinforced plastic and are stabilized much the same as are the skins in blades with continuous aft sections. Among the advantages of segmented skins is the ability to replace individual segments in the eveat of local damage. Because the skins are. in a sense, nonstructural, considerable damage can be sustained without destroying the basic structural integrity of the rotor blade. Also, with'this configuration it is easier to achieve blade bending stiffnesse of the values required for the natural frequencies desired in an articulated system. One of the greatest disadvantages in segmented aft sections is the increased difficulty in preventing water from entering the rotor blade. The number of segments may vary from 8 to 20 or more, and each joint between segments must be sealed.
is necessary to increasethe blade thickness to achieve sufficiently high section modulus and bearing arm. Commonly, this is accomplis;'ed by bonding metal laminates external to the upper and lower surfaces of the blade, and then adding a relatively heavy remntion or grip plate external to the stack of laminates. The retention plate contains the main holo(s), which may pus through the blade envelope or through top and bottom lugs that are extensions of the retention plates. Where the bolts pass through the blade envelope, it is reinforocd with Internal, metal filler blocks that effectively create a solid airfoil section in that region. When a tubular steel spar is wed, it may be extended inboard of the blade envelope and be fitted with a socket, or cuff, which is either damped or threaded onto the heavy root end of the spar. 1he socket may contain a single retention hole or two holes, d.pending upon the location and configuration of the lead-lag hinge of the hub. Here, again, the holes are through lugs that are an integral part of the socket and mate with similar lugs on the hub. 5-6.2.4
Tip Closures and Hardware
JiUIaUst
5.6.2.3 Wraparoad Skins A special form of continuous aft section is that in which the skin of the rotor blade wraps completely around the nose radius, providing both the upper and lower aitrfoil surfac.s in one piece. This method of contrtL crion may be used with any of the previously described internal stabilizing or strengthening members, although generally it i used with a solid extruded spar or a formed-section tubular spar. Such a skin usually is made of iluminum, although the use of other light alloys or fiber-reinforced plastic is not preclded. I he method of manufacture normaidy is to form only the nose radius in the center of the skin material, and to depend upon the spar and/or other internal members to control the remainder of the airfoil contour. A disadvantage of wraparound skins is the difficulty of maintaining cose contour :olerances, particularly in nonsymmetrical airfoils. Also, in order to maintain :he required weight and balacce, the skin normally is too thin to afford protection against erosion of the nose and, therefore, an additional erosion shield is required, 54.23 nreo End Retmade Root end retentions vary considerably from one blade design to another, depending upon the type of rotor system and the type of blade construction. The main retention bolts or pin(s) provide the interface between the rotor blade and the hub. Because of the high bending and centrifugal loads at this interface, it 3-44
rai1
Ai.flji
buI%4WA- hiIUvc
sUaa
tyfjMof
IASflA
and adjustable weights within the envelope at the tip end. Fixed weights are employed to provide adt quate rotor inertia, to control flapwite bending fmequencies, and to place the static chordwise CG and the nominal dynamic axis in the proper location. Generally, adjustable weights are installed in pairs, displaced equally forward and aft of the deon dynamic axis of the blade. These are used to equalize the spanwis' mass moment of one blade against another or against a master, correcting for manufacturing tolerances in weight, and also to provide a forward or aft adjustment of ti - dynamic axis to achieve equal pitching moments from one blade to another. With adequate precision in the tooling and methods of manufacture, the uced for either or both of the.:e adjustments may be eliminated. Tip closures may be simple flat plates or relatively complicated hollow, airfoil-shaped, monocoque shells, usually screwed or riveted to the blade envolope. Various sOapes are in use: some simply rounded at the end, some made in the shape of a wedge, and some with very unconventional planforms. Most tip plates or caps have a small protuberance at the extreme tip to facilitate flag tracking of the rotor. 54.2.5 Trim TrOs Rotor blades generally are fitted with ground-adjustable trim tabs. The tab may be an extension of the skin or of the trailing edal fioler strip beyond the nominal trailing edge of the airfoil, and may extend
for Aw4c panof
dop of O bbmk A
cow
ialty of 75% span For hums trad a& trim tabs am a&*u d by beafd tm upwald or downw..d a nmmusy to oqualisa the pifteba acemt chanac t9lst• of th* adibuAl Nbae. "54.A Teif
S
egd
c mmme Wok SMl
k
of motfmt
aOW a . To
e be to .bm dt thmu usihsm hdl ,,ain m .ul Si=h , it is duirbli that ah s # theospiwise nume Adbam1Mes be pemihi wkho~aihebie the chonhwiss CO. For eaa Ks.a pmawi, allnode weWigt f kaproperly locate Inthe bade owy correct a flapp*n natural frequecaq, but may chasg ouple I
Many forms of tuning weights am ued internally at locations alog the span ofthe blade. lly arevarious referred to a oantinodes l weights because they
a torsional fsqIency sao as to cause it to
ar placed at the point of maximun deflection amplitude of The the blade a itisvibrates in various harmonic mode*. purpose to chang the natur-al froquency of the ilade to avoid resonance with any
fications that can senousy liit the lifo of the rotor blade; vibrations are not transmitted to the irath but if the oiinnyntbeppttttl o'cupant..
posbible forcing frequencies, particularly rotational speed. The weigts may be bonded. riveted, or boltad to internal structur J members of the bladr, or may be suspended at the end of a cable, strap, or rod that is retained st the root end. The latter method of retention precludes high local stresses in the basic blade structure due either t. holM, or to contrifugal force because of the concentrated mass. It also pr"": M a "MW%,%"..... Aij :. SJ 4_Ac1 .,1i""hy bending, and permits the weight to be made of highdensity, nonstructural metal.
This control of natural frequencies is equally i•.portait in avoiding excessive vibration of the aircraft and high loads in the control system. Whether or not the rotor blade vibrations will be transmitted to the fixed system is dependent upon the mode of vibration relative to the nua Ser of blades .inthe rotor. Thus, it is necessary to ,onsider the entire systen, when designing a rotor blade for optimum natural
5-..7 Dedge ReWqiamb Regardless of the method of construction of a rotor bhlde, the detail design and the selection and distribution of matial muat satisfy a number of independent and interrelated requirements. The blade geometry Naving boen established, as discussed in par. 54.1, additional major considerations are strength, vibration, weight, mass moment of inertia, serviceabwiity, and cost. As a rule, a rotor blade that is des.ned to have a reasonable life under the applicable fatigue loading conditions will be structurally adequate for any static conditions. Therefore, major emphasis must be placed on design features that reduce the alternating strewn and make the structure as insnitive as posi ble to those rtesses. Alternating stresses are induced by response of the bledes to the periodic airloads, which, in turn, are affecte by the blade motion. The blade response it dependent almost entirely upon the mas and stiffnes distributions. It is exemnely important that thes distributions be such as to avoid any bending or t risional natural frequencits that are near resonance v Ahtany forcing functions (ace par. S4.2). Thi blade vibration frequencies may be broken ,own into flapwise, chordwiw, -and torsional frequencies, thse may couple together unfavorably to
------
strongly with a chordwlse bending frequency. Su'-
.
coupling often reults in high-frequency streW ampli-
In spite of all efforts to avoid amplifications of bending moments by control of natural frequencies, alternating stresses always will exist. It is of prime irportance, therefore, that the detail design mimiz,, the tolerance of a rotor blade :o these stresses. Materials selected, whether metallic or nonmetallic. must be capable of providing high fatigue strength. To this end, any form of stress raiser - e.g., notch, hole, or sudden change of section - must be avoided in areas of even relatively low alternating streUs. Techniques have been developed that now make wcldi-ag a vm~bi• mfu(hd fUo fabliitfion of rotor blades. However, care must be exercised in the placemeat of the weld, and adequate quality control over the process must be assured. Holes in areas of high stress also can be avoidc; 1hrough the use of adhesive bonding. When the joints are designed with care, stress concentrations virtually can bc eliminated. Bondcd joints albo act as a barrier to the propagation of a crack from one structural member to anothe'. In all of the blade confriuretions discussed in the earlier parts of this paragraph, adhesive bonding generally is the principal method of joining. 54.8 Toe11" NW QAltY Cer" RqdremetTwo principal catgories of tooling for the constraction of rotor blades are the tools for fabricating the main emponewtp and thow for assembling the blade. In the case of molded flber-reinforced-plastlc blades, these may be combined, and the spar, skins,
"'-:-
I
J
.9-
"'.N
'
\_c.also may be ma.'. .... .. mbb eool. Dies for the manufacture of extruded aluminum spars are relatively inexpnasive; however, the machining and other operations involved are likely to offist this cost advantage if the spar is tapered in any way. Frequently, it is difficult to maintain the required tolerances in aluminum extrusions. Formed sheet-metal s~ars and shear webs, or longitudinal stiffeners, usually are made in a multistage roll forming mill if the blade is of constant seetion. T -oling is more expensive than extrusion dies, but is more durable and produces parts to very clo•e tolerances. For tapered blades, it is necessary to stretch-fonr the parts. Tools and capital equipment for this operation can be quite costly, but very close tolerances can be held. Tubular metal spars may be made by any of several nmethods, all of which are some form of swaging Tooling costs generally are quite high. Quality hazards associated with these proceed, include mandrel pickup and the enlargement of otherwisc nqgugible or easily removable metal defects. Adhesive bonding requires large, specialized tools capable of applying accuratcly cunitrolicd heat and pressure while maintaining close dimensional tolerances. There tools may bz "unitized"; i.e., they may contain built-in sources of heat ang pressure. Htat may come from the electrical resistance "calrod" type of inserts or heating blankets, or may be provided by steam or hot oil passages. Pressure sometimes is applied through pneunatic cells contained in the fixture. Unitized tools have the advantages of being semiportabic and of bcinj capable of providing differentf values of temperature and pressure in df.ee_t 7,nn~ Mn,na- reviioi.-A~ ._•i._• for f ths_r me i nf of-,tn•... _. mutriouil .---different
Rcqardims of the typo of tooting, prcci oilnti 3ls are required to assure that proper heat and pressure have been applied. Printed chart records are desirable, end provisions should be mado for the proceasing of samples representative of each individual blade assembly that can be tested to destruction. Even after having maintained such control, it is desirable that some form of nondestructive testing (NDT) be applied to the final assembly. The most prevalent NDT method is the ultrasonic scan, which reveals unbonded or poorly bonded joints. 5-6.3 BLADE BALANCE AND TRACK Individual production rotor blades must have both dynamically and aerodynamically similar characteristics. Dynamic similarity is achieved through maintenance of a specific mass balance by the addi. )n or removal of weights on the blade. Aerodynamic similarity is achieved by maintaining close airfoil and geometric control, or by adjustments, such as with a trim tab. Determination and confirmation of dynamic and aerodynamic similarity are accomplishd by phyt'.Allwy bh.nAng anid tra-.ring ,,ah blade against a master blade or set of master blades. 5-6.3.1 Effect of Dedgn Considerations in obtaining dynamically and seaodynamically similar blades must begin with the design. The selection of materials and the construction of the blade should be made with interchangesbility as an ultimate objective. In making material selections, trade-off such as sheet stock versus extruded or forged muterial must be made. Generally, a weight advantage can be realized by tie use of sheet stocik. llhwcvcr. the shect siock ~ec.te the flori~oig of'fteacLM '--
"'
and the type of joint in each particular zone. A disadvanlage is that each rotor blade type or subasnernbly requires a completely new tool with heat, pressure, and cooling provisions and relatively comiplex controls. The other principal assembly method is the autoclave. Both heat and pressure are provide6 by this piece of capital eqisipment, and the tools that hold the blade components and maintain dimensions during bonding are relatively less expensive than comparable unitized tools. However, unless special provisions are made, all areas of the blade receive the same heat and pressure. This can be a distinrt disad-
may not produce a close-tolerancc airfoil shape. This type of trade-off procedure should be followed for all major components of the blade to insure acceptable balance and track and ultimately, the interchangeability of each blade with other blades of that specific configuration. Bec:ause the control of the weight of individual parts within a close tolerance could result in extremely high costs, some adjustment of the weight of the blade must be provided. This adjustment should permit the addition or removal of weight at the blade tip and, possibly, at the blade root as well. In many blade designs, the adjustable tip weights are installed on at
vantage since more heat input is desirable in a region such as at the root end, where there is considerably mom mass, than in a light sction of the blade. A hybrid method of assembly employs a tool that contains its own pressure source, such as pneumatic cells, but that is placed in an oven for heating. 5.46
least two separate chordwise attachment points (par. 5-6.2.4). The location of thc adjustable weights at the tip takes advantage of the large balance nrm about the reference datum, which usually is the center of rotation. Spanwise balance is achieved by adjusting the total weight at both attachments, whereas the
AMCP 706-202 chordwise CO is corrected by transferring weights between the chordwise positions. The limits of adjustment are reached when either attachment is completely empty or is completely flled with weights. To establiah individual dynamic balance, !%oth the spanwise and product moments must be controlled to maintain a common dynamic axis for all blades of a particular model. The dynamic axis X!is expressed as follows: it J xydm X ,in. (5-8) NOR! where dm - increment of blade mass, slug -= location of flapping hinge from the center of rotation, in. m, " mass of spanwise increment at inboard end of blade (e). slug nmk - mass of spanwise increment at outboard end of blade (R), slug R - blade radius, in. x
= chordwise distance from blade leading
edge to centroid of mass increment, in. y - spaunwise distanct f(oun flapping hinsc to centroid of mass increment, in. Eq. 5-8 implies that the weight of each element or component must be rigidly controlled. However, in practice this is not necessary because a system can be established to match rclativ'tly heavy parts with those that are on the light side of the tolerance scale. Because it would be quite cumbersome to match or select each and every part of the blade assembly, only those components that make up the bulk of the weight need be considered. This method of selective assembly divides the rotor blade intt, four main camponents or groups: the spar or spar assembly; the leading edge material, including ballast; the aft acction skins and stabilizing material; ared the trailing edge reinforcement. These four major components are selected bemuse they comprise the basic structure of a blade and extend the full length of the blade span. Variations in the weights of t'e remaining parts have little si3nificanoe in the total weight and balance of the complete blade. 54-.3.2 Comneeet Lunit Weigtts By selecting the major components on the basis of their respective weights and moments relative to the Uaveilcble weight adjustments, virtually all blades can be balanced to a master balance blade. The weight -variation in each part shal! be determined by the available capacity of the attachments for adjustable weight. Weight limts for each part may be calculat"ed by assuming that all other parts are of nominal
weight and that one of the weight attachments is completely empty or full. When two tip weight attachnients are used. the acceptable weight tolerance on forward components (e.g., spar and abrasion strip) is lim.ited by the forward tip weight capacity, whereas the weight tolerance on aft components (e.g., trailing edge and skins) is limited by the aft tip weight capacity. When this method of weight adjustment is used, the limit weight for each component may be obtained by the solution of simple pairs cf simultaneous equations. The equations are iet up in terms of spanwise and chordwisc (or product) moments where the sims of the moments of empty or full attachments and the two unknown weights arc equated to the sumns of the nominal ,nomente on the same compon.nts, aF shown in Table 5-2. The steps that follow (using data from Table 5-2) show the solution for the minimum and maximum weight for one part (an abrasion strip): 1. Minimum al'owablc weight (forward attachment assumed full):
a. Spanwise moment is: 0.19(155.50) + 82.6. WJ.., + 155.50W, 459.30 + 15.55 + 15.55 b. Product moment is: 0.19(155.50) (0.75) + (82.61)(0.558)Wjm,, + (155.50) (2.35) W. - 256.30 + 11.66 + 36.54 c. Sol-.ing these two equations will give W,,•. the minimum allowable weight for the abrasion strip. lb, if all other componcnts remain at nominal weight 2. Maximum allwwable weight (forward attachment assumed empty): a. Spanwise momeqt is: n c '+ 82.41 w + _L15.50 ul 459.0.+ 15.55 + 13.55 b. Product moment As: 0.0(155.50) (0.75) + (82.61) (0.558)WIR• + (155.5) (2.35W1 ) - 256.30 + 1.1.66 + 36.54 c. Solving these two equations will give W-1 ... the maximum allowabhl weight for the abrasion strip, if all other components remain at nominal weight. In both the solutions the valwts W, and WJ, the aft and forward adjustable weight•, respectively, must be :50.19 lb. the maximum capacity of the adjustable weight attachment. A sind;ar set of simultaneous equations is solved for each of the other three critical components. Nomograms can be prepsred for convenience in the selection of the four oi more critical weight components. These nomograms combine into a single graphical format all the mininim and maxitpum c&'imponent weights deterwii~ed by the procedure stated previously. Similarly, th.se results can be com-
TABLE 5-2. EXAMPLE OF NOMINAL WEIGHT AND CG LOCATIONS WEIGHT, lb
PART
A
____________
ABRIASION STRIP
5.56
MOMENTCG LOAT ONS PRODUCT, SPAN WISE. in. CHOfIDWiSE i. -VSA-NWISE, Ib-in.'A x B :D Ib-In. 2C x D-* c B 266.30 459.30 0558 82.61
1 AJUSTABLE
FORWARD'
0.10
155.50
0.750
15.655
TIP WIHS
AFT'
0.10
155.50
2.350
156.5
36.54
6
'MAXIMUM CAPACITY 0.19 lb
PART
SPANWISE
SPANWISE
PRODUCT
PRODUCT
WEIGHT,
MOMENT COEFFICIENT,
MOMENT, lb-in.
MOMENT COEFFICIENI,
MOMENT, lb-in.
A
B
E
SUB-TOTAL COMPLETED BLADE UNBALANCED216939. ADUTBE
TIP WEIGHTS ADDED
00984
FORWARD
.-
IAI z
jI
155.50
I r l13654
'
SUB-TOTAL COLUMN E SUJB-TOTAL COLUMN C
DYNAMIC
FORWARD AFT
2.
35406.
-Mb.0051 -
0.005
155.50 155.50
TOTAL FOR TEETER
3671.2 2196.3
lo
-0.8 -0.8
().7in -
-0.6 -1.8
116.60 365.40
21473668.8
BALANCED BLADE
FINAL DYNAMIC AXIS (CHECK)
9.8
21963_371.
AXIS (CHECK) ADJUSTABLE TIP WEIGHTS
116.60
216.\~*~
S1JB-TOTAL PRELIMINARY BLADE *BALANCE *PRELIMINARY
13.1
___
TOTAL OF COLUMN E TOTAL OF COLUMN C
3668.8 2194.7
167in
NOTES: (D THIS WE.-GHT ADJUSTMENT IS MADE WHEN THE BLADE IS TEETER BALANCED. THF NEGATIVE SIGN INDICATES WEIGHT WAS REMOVED. Q~ DYNAMIC AXIS AS MEASURED FROM THE LEADING EDGE.
AMCP 706-202 bined into a system employing a digital computer to
closely controlled weight and balance system during
provide rapid component selection from a number of random-wcight parts,
manufactufe. The prerelcac tracking of all other blades will be made against mastcr tracking blades.
For final balance the spanwise and product
The master blade(s) are blade(s) that have been fabri-
moments of all of the blade components, including th,. paint and adhesive, are obtained for the unbalanced blade. Again, two simultaneous equations can be written. The unbalanced spanwise moment and the forward and aft tip weight moments should be equated to the rt juired spanwise moment of the master blade. The unbalanced product moment along with the forward and aft tip weight moments should be equated to the desired product moment. Solving thesc equations simriltancously will yield the additional weight required at each location for dynamic balance. A summary of the balance procedure is shown in Table 5-3. By establishing a weight tolerance for each of the major selective components and using a consistent method of part selection, the blade assembly will, in nearly all cases, balance within the capacity of adjustable weight attachments. Upon final assembly of the blade, it shall be balanct,-checked againzt a master blade. The tolerance on the actual balance depends
cated as closely as possible to design specifications and to as precise tolerances. These master tracking blades are produced so that, when they are installed, the controls arc adjusted to the nominal position. It then can be ascertained how much deviation ot toler. ance may be allowed on production blades. Consideration of allowable tolerances shall includc the crew comfort levels defined in MIL-H-8501. Interchangeability with mauter blades must be determined either on a tiedown aircraft or on a suitable towcr prior to relesou for random installation. At least one master blade must be tracked with each group of production blades. The blades should be tracked at several rotor speed settings typical of those that will be encountered during operation and at several values of collective pitch, with rotor speed held constant. Track readings shall be taken for each blade at each speed and pitch setting. Typical data are shown in Figs. 5-29 and 5-30.
-oil Woil iu iy; of' uzolt.C.g., fully itiuiulait
UK
hingeless, and the size of the blade. Tolerances of the order of 10 in.-oz are not uncommon. This physical balance of the blade must be performed on a balance stand capable of registering the blade spanwisc moment to within the specified tolerance. Weight should be added or removed as required to balance the new blade. The balance master shall be established as that blade to which all other blades of a particular part number or series shall be balanced. This demonstra-
-
tion shall be accomplished by balancing each blade
either directly against the master balance blade or against a calibrated mass balance for which the master blade was the calibration, stanrakd. The spanwise teeter balance discussed previously demonstrates only that the blade will be in flywheel balance; however, the dynamic chordwise balance still may be out of tolerance. Dynamic chordwis. balance, therefore, must be checkcd by tracking the blades at various rpm and collective pitch settings. If the selection of parts was controlled during the fabri-
cation of the blades, a minor adjustment, such as
7
3i.0_ S
BaADE
o
--
_
c 1
0.5
1.0 1.5
....
200
-T... 210
. 22W HUIQ4 WEED. rpm
230
240
Figure 5-29. Track With Varying rpm (Zero Collective Pitch) 1.5
.o0.15 "
B-AA
--
5 0
moving adjustable tip weights forward or aft, will correct any dynamic chordwise deviation.
0o..
"J3.33 Track "To confirm interchangeability, each blade should
1.5
be tracked prior to its release for installation. However, tracking of hir4eless blades is difficult because "the deflections at the tip arc small. Interchange"ability of these blades can be confirmed by using a
-
1.b
--
4
COLLECTIVE PITCH
Fipure 5-3,
0 0.
Track With Varying Collectbe Pitch (Comntat Rotor rpm) 5.49
A typical plot (Fig. 5-29) cf the tracking data r- orded by any one of several tracking methods during an rpm sweep of an articulated rotor indicates
ment has much influence in the establishment of the minimum mass distribution for the rotor blade. No weight savings can be realized beyond the limit ir-
that blade (A) is aerodynamically similar to the master; however, an incidence or pitch adjustment is required to correct the blade track to zero. Blades (B) and (C) can be corrected by downward trim tab adjustment. The appropriate adjustments should be made to bring all blade tracks within the tolerance level compatible with crew comfort levels previously established. This tolerance will depend on rotor size but commonly will be equivalent to dIfferential blade coning angles of the order of 5 min. A typical plot (Fig. 5-30) of track data during a collective pitch swoep with the same articulated rotor depicts blades (A) and (C) out of track d.ic to dynamic unbalance. This condition is corrrected by moving a portion of the adjustable weight of blade (A) forward while that of blade (C) is moved aft. Blade (B) i: seen to be dynamically similar to the master blade without adjustment. Weight adjustments and/or trim tab or trailing edge adjustments rmvide hlnd -"will that are dynamicaly and aerodynamically alike, permitting interchangeability with all other blades of that configuration. In both Figs. 5-29 and 5-30, the reference master blade is shown as a horizontal straight line through zero without slope, Several methods of tracking blades may be employed; the accuracy, safety, and reliability of the electronic trackers provide excellent results. In addition to tata flat and collective tracking data, the blade pitching moments should be determined with a suitable calibrated load cell. Pitch link forces should match the master blade pitching moment
pGed by this requirement; thus, a point exigts beyond which an increase in the strength-to-weight ratio of the material cmnnot reduce blade weight. Although a high strength-to-weight ratio is desirable, more important factors are the ratios of both the fatigue strength and density to the modulus of elasticity of the nmaterial. The tetal loads on a rotor blade cannot be predicted by a straight forward examination of rotor thrust and centrifugal force.rTe loads depend upon the response of the blades to the periodic airloads which themselves are affected by the blade motion. The blade response also depends heavily upon the mass distribution. A change in stiffness affects the bending moments, deflection, and radii of curvature of the blade to the extent that the response is changed. It is impos.ible to predict - without a reevaluation of the blade response -. whether a change in stiffness will increase, decrease, or have no effect on the radius of curvature of the blade. In other words, the radius of curvature of a rator bWade do, %A not have the simple proportional relationship to stiffness that exists in a static structure because the bending moment is a dependent variable. Nevertheless, the following equation from simple beam theoty for the radius of curvature r is applicabtle
demonstrated that vibratory levels do not exceed the limits of MIL-H-8501 and that life-limiting oscillatory stresses are not induced. $&4
ROTOR BLADE MATERIALS
As discu a-d in par. 5-6.2, a relatively broad variety of materials may be used in rotor blade construction. This piiragraph considers the major factors that lead lo the selection of specific materials, based upon the inherent properties of the materials and irrespective of the details of construction. Helicopter rotor blades arm unique in that many conditions that must be met depend upon various combinations of material properties. A rotor blade must be designed as an integrated part of the complete rotor system. One specific requiremenW is that the mass moment of inertia of the rotor system must be of at least a minimum value to provide sntisfactory autorotational characteristics. This require-
5-s0
=
-
-
Zt--I
.
(5-9)
..>
..
O
and bending stress crt in a particular material with modulus E is or's
psi
(5-10) AE-
Substituting Eq. 5-9 in Eq. 5-10, _.-,p r where c - distance from beam neutral axis to outer fiber, il. E - modulus ofclasticity. ps 4 I - moment of inertia, in. M - bending moment, in.-lb If a stiffness change is made in such a way that the distribution of mass and stiffness is unchanged, the blade response, and thus the radius of curvature, also wid be sub ntantially unchanged. Then, as in Eq. 5-11, the blad. bending stress will increase in direct proportion to the material modulus of elasticity E. It follows that the most desirable rotor blade material is
.
---
•,
) -
__
-
--
"
_ '_
-
the one that has the highest ratio of strength to modulus of elasticity. Any material with a high modulus of elasticity that does not have a proportionately high strength is undesirable. Table 5-4 is a comparison of the ratios of material fatigue allowable (FA) to modulus of elasticity E for a sample of available rotor blade materisIs. The comparison uses fatigue strength because this factor is of primary importanoc in rotor blades. The values given are based on experience with actual structu.res and arc less than the values obtained from laboratory specimen data; however, they are presented here for illustrative purposes only. Syccific values of fatigue strength for metals, plastics, and sandwich structures are contained in MIL-HDBK-5, -17, and -23, respcetively. Care must be exercised in using aniy given values for fatigue strength since the configuration of the specific component as wull as the necessary manufacturing processes may adversely affect the material properties. Considering only ratio FA/E as the criterion, Column 3 of Table 5-4 indicates steel is suptrior to aluminum, and Fiber.glas or graphite is superior to
material is used in a rotor biadc, other considerations are necessary since a rotor blade operates in a rotating field. In this condition, strrin compatibility determined by the ratio of modulus of elasticity E to mass density p becomes an important factor. Ire a rotating field, the centrifugal force (CF) generated by each blade element is proportional to the mass density of the, specific material and the position of the element along the blade radius, or span. When two continuous spanwise members, each of a different material, are side-by-side in a common centrifugal field, each will tend to strain an amount that is proportional to its respective mass density p and inversely proportional to its modulus of elasticity E. In most cases, the two members are bonded together with an adhesive that can transfer load from one to the other by shear, causing them to strain equally. This bcing the case, the material with the higher value of -he ratio E/p will pick up load from the other material and bL strmsed higher than if it were rotatring bv itsclf. Theis, it is drgirable tt6at two or more materials, used in coniinetion. have fairly similar valuws of E/p. Column 5 of Table 5-4 indicates that
either mptal
aL!umniwnm and at"_. a
Rnrnn ir not nkrtifir!i"rI
nttrtivr.
Wood (spruck) i, highly fati~uc-resistar:t, but also has
disadvantages that preclude serious cohsidcration for prescri-gencration helicopters. Up to this point, the discussion of materials has dealt with bludc b(.ndirg only. When more tian one
I4
TABLE 5-4.
MATERIAL ALLOY STEEL ALUMINUM -'E" GLASS/EPOXY tUNDIRECYIONAI. BIDIRECTIONAL "S"GL.ASS!EPOXY LI•NDIRECT!ONAL BIDIRECTIONAL BORON/EPOXY UNDIRECTIONAL BIDIRECTIONAL GRAPHITE/CEPOXY
cluminum is a~ceptable. Boror and graphite ve compatible with caci other, but cithbs
should be tsed
with caution ir combinition with steel, aluminum, or Fib;rglas.
COMPARISON OF MATERIAL PROPERTIES
-"1
"".-
101
ih tchi respc.t, and Fibe'g!as in combinaioun with st". or -rr.nn_.tibe
E, 29 10 6 3.5
FATIGUE
FA
FATLGUE ALLOWABLE
FA
-_
± K0.00 6,000
0.0010
__N0 0
E r. P. DENSI Ibin 0.28 0.
Pn
103
8,700 4,200
0.0014 0.0012
0.065 0.065
92 54
8 5
t 9,700 t 4,900
0.0012 0.0010
0.074 0.074
108 67
36 21'
t 26,000 413,000
0.0007 0.0006
0.074 0.074
486 284
4
+
-:;•
".
UNODIRECTIONAL 30 ±40,000' BIL!R- .TI NAL 18±20,000' SPRUCE 1.4 ± 2,000 DATA EXTRAPOLATED AND.'OR ESTIMATED FROM
0.0013 0.053 0.0011 0.053 0.0014 "NUMIERJUS SOURCES.
S65 340) 88
S-
In vimw of Sr;wina prmsures to use advanced cornposit=. in aarcraft structures, it is appropriate to examine the beciefits, if any, to be dcrivei from t~teir application to rotor bladm. The outseanding attractions of such materisis arc very high stiffness, high strength, end low weight. It has been thown that ft iratios of these properties - rather that, t&;,absolute values - art of primet importance. From the standpoint of fatig-.ie resistance (Columnn 2. Table, 5-4), these matcri-is appear to be very compatible with mrore conv.-ntional materials for use in rotor blades. The que~stion, then, becomes whet",,r there arz overriding advantages to be gained from other' chai actcristics, such a3 ballazflic tolcraro=, or the high value of the ratio Elp. Tht rotor blade dynamic repcinsce is highly clcpcndent upon the rotating natural firoquencies of the blade, and it is necesary that the blade be designed to avoitht frquecie ae inresnane wth ny ortot natural frequenc expression The functions. cing cy of a rotating beam wR. is K-O
which is wi, o W
73.6 27.1 272
Now, assuning that the stiffiles El is increased to 40 10' lb-in.', 1 ' R, =%-1390~ 47T20 -%46F1_0 78.2 rad/sec Thus. for the iirst flap-wise bending mode on a typical
hiaiged rotor blade, a 100% stiffness change iesults in a chznge in rotating frequency P_. wRi 4.6 rad/sec, or 6%. This mode of vibration is crit~ical in an articuelated rotor, and *he only effective way to centrol it is by varying the mnaws distribution, because little can be done by changing the stiffness. Further examination would show that the first F77~TYchoidwise mode of vibration, as well as the hiiher +1 A!L rdse Qt) zL in hnth planes; is affiected significantly by lei 0,+mt'jdse s-2 blade st~ihress. I-cr hingeless or semirigid rutors, all modes are aflTecled significantly by blade stiffness. In( whrethese, typcs of rotors, the higher mniodes of vibration a., co.afficient which is dependent upon m$ss are m~anifest primarily in blade: stress levels, as opanJ stiffness distribution and has a daiTcrposed to vehicle vibrations. F~or multibladed rotors, c-it value for each mode of vibration, dithe higher vibration modes also can contribute signiEl= stioTn'ss.lbi.ficantly to vibratioirs. E co=flicientss deenet po as dsiu Material selection also can be very important in t ofiien a dtepmoden t ofvbrtonmssdimestionrotor blade fabrication. it now is possible to produce ti~ hldf vbraion te mde diensonnounif'orm bladc cross sections in any of the les3 available materialb, although it genci allly is easier MI!5~~~~~. em suz/n ".rd~i Ap!1rgtha in lActh h- mnA inn.n I - lengthoa pe, rads. stanices whert noncompatible values oAfLp can cause Q -taioalsee, adse:high stresses in flight, it also is important to avoid For vaucs inge,,beas. ad K, tc forthe irs triteril cobinaionswidifferin dfferng cJithrc mod ar: cints f terma expnsin. Sch emblics ,zan a, a., -~50.0. a3 - 10-5.0 develop high rtsidual stresses as a result of adhesive K, - 6.38fsKz = 17.65. K3 - .0bonding operations. To examine the effect of stiffness El on natural freIdeally, a rotor blade should be made of materials uiuency, an example is presented: a constant-crossthat are highly resistant to both corrosion and crosection blade of 25-ft radius with a weight of 4.0 lbý/ft sion. Corrosion resistance of the nonmetallic cornand aflapwiseElof20 x " lb-in.' A tip speed of pstsi ihyatatv n :r eifunili 680 fps is assumed, giving a rotational speed 01 of 27.2 psieishglatrcvendarbeifetaln selection. In a monolithic composite blade, encematerial rad/ec; rad/ec; ence it is necessary to protect the forward portion against erosion. The most effective materiali for thiz purpose 20 X (1,Y 4'R (25T l2) + 6.38(27.2) are stainless steel, niickel, cr cobalt abrasion mseum~d. 4 12Of X(2 the elastomeric materials, tht urethtane arm 12 )(l2 ) 32.'~superior and are very durable when sub~tetd to sand, but generally have been found to have short --%69_5_7__ ,~- j-4_ liv~es when rain is a significant part of the environ73.6 rad/sec ment. _______
______
-
-
It has been determined that rotor blades are vulncrable especially to lightning. To avoid damage to blades oubject to lightning strikes, provision must be made for low-resistance paths for the high currents that are characteristic of lightning. The basic lightning pro:cction requirements for all acrospac4 systems are given by MIL-B-508'. For blades constructed cf composite materials - inherently poor conductors - or even those of all-metal bon ded con-
Critical helicopter components are sub;e.t to a load spectrum characterized by a relatively high-frequency oscillating load content. Characteristically, the rotor system - particularly the main rotor produces and endures the hi3hest cyclic loading. A fundamcntal design requiremont is long life of rotor system components. Resor.ant conditions thai proV1agho dn ua11wnig Bywu..levCes within a stcss contponrnt must be avoided. However, compliance with these requirements can be verified only through the correlation of flight test and component fatigue test data. The discussion that follows supplements the
ditions, can define a preliminary load spectrum, consisting of load magnitude and frequency, as well as frequency of occurrence. Thes data, when combined with section property and theroctical stress concentration factors foi a component design, cast be con"Yerted to steady and oscillatory S-N (stress versus number of cycles) data. Therefore, preliminary component life can be determined based upoa cumu!ative Oaniage and notched and unnotched material or similar fatigue test data. Coupon fatigue test data must be used with cure aince these data usually will not reflect accurately the effects of manufacturing processes that aic peculiar to a specific component design. S-A' test data for components of similar design and manufacturing process are more useful in the preliminary detcrmination of component life (see pa;. 4-!, AMCP 706-201). Although the previously described method can be employed in the prelimihary design phase to predict component life, a mote rigt•'ous analysis of :ompor.ent fatigue and flight test data must be performed to
description of fatiguL life deterinnnation given in par. 4-11, AMCP 706-201. Corrosion has a rapidly degrading effect or the fatigue strength, and related life, of a particular component and this effect is difficult to predict in an ac-
determ;ne the final life of the component. In the laboratory fatigue test it is necessary to simulate the actual combined loading conditions, particularly in areas of local attachment or where actual load paths may be in question. For example, tkt r ,,•/r blade
curate quantitative manner. Corrosion-resistant ma-
root-to-hub attachment, a meanie.1f:
terials ani/or proven corrosion prot•,ction methods
tion of the fliaht condition ia¢hnokc
thus should be used to cbviate the necessity of con-
flap and chordwise momaents and shears pa vaixtio
sidering corrosion in the daternn.iation of rotor system component fatigue life, Fretting is the erosive failure rf the metal surface ias the result of small displacernents of heavily loaded
superimposed or the centrifugal force. Because, 'I is not feasible in many caste to include all associated -Nr influencing components in the fatigue test, it often is important to simulate loca: flexibilities offered by
mating parts. All preventive methods practicable should :e employed in the dasign and development
flexures, bearings, etc., or to simulate local force inputs to the test article such as those offered by pivot
phase to preclude the occurrence of fretting betwe.n components, particularly crifical, highly stressed blado/hub retention areas. Fretting occurs commonly in areas such as tenwon-torsion strip packs, bearings, (particularly low-angle oscillating applica-
point friction or lead-lag dampers. The failure data acquired for an assembly quite often will involve the failhre of only one component of that assembly IThis component then becomc the limiting factor in the life of the a&scrmbly. If such a
tions), i•.vtntion hole bushings, and blade/hub attachment fittings. The degree of sucrxss in pre"venting the occurrence of fretting is determined
component is a replaceable item, it can be replac"d periodically during testing, as failures occur, in order to acquire failure date for the longer-life conipo-
through careful inspection of various components
nents. Although the individual components o! the
that have been subjected to fatigue tests thnt simulate actual installations and loads,
assemhly can be tested separately under simulated loading conditions. 'esting of t~e complete assembly
structio., adequate protection against ightning damage shall be demonstrate4 by test (see par. 8-9.4, AMCP 706-203.) Valuable preventive design guidelines are given in Chapter 7, AFSC DH 1-4.
5-7 5-7.1
Elumi.
)
J
Rotor system components with long lives can be attained by implementing a combination of techniques in the initial design. Preliminary calculations of rotor system natural frequencies and loads can be made with a reasonable degree of accuracy for a prescribed number of representative vehicle flight conditions. The mission profile specified for the vehicle, coupled with load calculations for spicific flight con-
ROTOR SYSTEM FATIGUE LIVES GENERAL
\~
ct•n-
.j]
.c
' I - '
1
radii in jowpgle or over similar fitting area, should be included in the specimens tested. Fatigue or endurance tests sAaI be conducted on a sufficient number of coupons for each cond~ition. The number of coupons nevcr shall ike ks than five. Howover, if the standard dcviation oftimt data points for any test condition exceeds 15% of-the mean stress, additional couponu shTI be tested. Coupon tests may be conducted on any suitable test fixture or stand capable or applying an alternating IDad. Thec alternadin,4 load may be superimposed on a steady, or mean, load to produce a load condition as showa in Fig. 3-31. In the normnal rotor 5-7.2 ENDURANCE LIMIT TESTING blade load sprztrufir, each load condition is a combi5-7.2.1 Getk nation of steady and alternatting, loads. Therefore, the Endurance limit testing generalily is required to obuse of this loading cirsdition for coupon testfng is for life. sc~vicc the to guarantee tas:- data adequate kocooimended. rotor blades (sec, par. 74.2.2.2, AMCP 706-203). This Coulson designi depcads upon the type of material is testing in which the material and/or part is subbeing tested. When the material is sheet metal stock jected to repeated cycle& of load, with or without a some forim of the "dog-bone" co~pon should be steady, or-constant, load maintained. 1'hcr zndurance; used. The transition from the gage sqciion to the grip limit for most haomogeneous, near. isotropic materials areii s/rai be stsch as to eliminate a stx.ma concentrahas been established, and is defincd in MIL-HDB'Ction due to section change. An acceptable radius for 5.. Fatigue data for plastics and sandwich construcsuch a tranisition is given in Ref. 47. Additionally, uon~~~ II-likNIIan ~ ~ ~ ~ arfrn5IlOa ~ ~~~Cl c scsc' be tak~en Wuit the S4-cdgcandiation. of sheet tively. In most instances, these data are prescnted for stock coupuns. The machin'ng of the edge should be both smooth end notchod spoci'fcits. However, with cotolds 'topenthraldaaonf siudbegpeparedo o nes ton pedgen.Edgesa coth oled so:i the introduction of many and v'arivd reirforced onoitionoye udessingtcnld bgriteaipared oihedmt~ra to plastics and advanced composites, applicable fatigue gri line drtor by burlishidcng.aditin. maingfne;base test rcsults aire not yet available, in the lillercturc. bat ipermidtitc e w th beeth~ishmner g n ca lurimhit bnc O advanrd such for required bk will Basic ters thur oc a eea-ihdta ilpri Ac ii materials. and safety faz.toi reductions from a reliable reference. It is always nemcessy to excrcisec care in u.-.ing data that are rclatcd to dic shape of matcrial under consideration. For' example, the. ume of tihetct stock in rotor bladr? design dicatetc the usc of teiision-tensioniATfN IC fatigue data, whci' avadahsb, to prodia! skin fatigue SRS life or thr life of an~y part made from sheell stock, P-owcvtr. it nmy be more app:opriaae tw use R. R../ M',ore rotating huam fatiguF, data when solid bar or plate stock is integieted into the blade. denign. Additionally, an appR.;Ople notc~i factor, either intirreat in the .*sig'i or reu!VoS frorn the manufacluring prtness is of prinary iw\portaftcc.MAIU To confirm a mateti sixetieýA. it miay be nc(..MXKU sary to coriduct cokipan tzsis to sub itanaate a perticular mnaterial condition trot covcred in the curreni STEACDY literature. In conducting thte ci.Nupoiý fatigue tests, it is &iHESS extremely impiortant that the tvit material hA& becr MIMN!MU1 subjected to the processe asico*iated with fabrice.STRESS usualy isi f~al n re oidd the eSlCects of load tr'adfer beween components. For bonded, wmlded. or odierwise permanently fastened assembhie; kndivikWua component test data must be acquired; or S-N dam~ of stich components in like material, piwocees, and configuration may be empiovad, if available. These data should be modIfied by use of the Goodman diagram or other acceptable means to reflect the presece of steady loads as appropriate. Methods of obtaining acceptable cornponent S-N datuiare discussed In par. 5-7.2.
ticn of the ctitiWu compor~ent. It.addiion, tests srhaillI
_
I I_ be corducted vq thr peltinent rnatvi all slsapc, using stock, bar, or plate as the dtzisi' dactttes. OtherT!?E akneeta of the propv~nd compoecnt configurati.'-n Flgw-e 5-31.. Alleeoath6g Stre Superimposed on such as edge condition, fillet radii, oi th~Mr7 bend Steady Stress 5-54
K.
)
~AMPk 706-202
Wtien ntocs* thickness pef mits in the case of plate, bar, castings, and forgings - it is preferable that the specimeris be machined with a round cross section. The circumnferentia surface should be pohshcd in order to remove any Ctress conoentralions or notch effccts. In order to predict allowable fatigue stre'igth tor a part or component, notched specimuins should be tested to determine the resulting fatigue strength reduction. Experience has shown that a typical value of this notch fa'ztor K., is about 3.0 fov almost all metals. Tht notch factor K7, which may 6e calcu. lated or be based on test data, isthe ratio of the peak stresses in notched and unnotched specimens. Data from tests conducted using the appropriate configuration of coupons will establish an S-ps' curve, In developing each S-N curve', a minimum of five specimens should be tested at varying alternating loads with the same steady load, The stress level at which no failure occurs after 10' cycles for ferrous metals establishes the endurance limit of the material. A family of S-N curves, each for a different steady load will provide sufficient data to 00"t a Go0dima- diagrarn I=e Par. _"-!1. ANMCP 7 201). The Goodman diagram, in turn, will permit the -e~g.,
convorsion oA a particular load condition into an equivalent load cori~ition of different steadly and alternating load levels. MIL-HDBK-5 presents the data in a constant-lire diagram rather than the Goodman diagram. in any case, a diagram constructed from coupon data should be revised as compoaznt and assmbly test data art generated. S-7.2.2 Noumetal The use of plastics reinforced with glass, gralihitt, and/or other advanced composite materials ;n thc construction of rotor blades will r-Auire the development of both S-N curves and rational Goodman diagrams. Additionally, the processing of these matcrials is subject to variations among manufacturers. Therefore, care must be exercised that test specimens are representative of the rnatenal and processes to be used in the blade construction. Particular attention should bc given to fiber orientation with rcspect to the principal axis of loading. The establishment of a family of S-N c~urves similar to Fig. 5-32 is an acceptable method of determining allowable fatigue strength for a particular reinforced viiutic-a
no-' adunn',j1 compoite Mat.rial (ACM).
These fatigue strengths, or endurance limits shall be
T
-
-.
CROSS PLOTTED FROM FIG 3-12, MIL-HDBK-17 FOR ..- MIL---77 POLYESTER RESIN
40
ztU 30
,.
0.20-
)
30LE
-h..
1ir~r5-3t.
A
~tng~~ssCycls a Vaiou Stedy tree Lvelst~0.300tt
4-_3
-
'
.
establishod for a particular steady load at a minirain 5 x 101 cycles without failure. 71-A obj'!Y-ivz of the coupon teW. is to establish or verify the endurance linit of a material for sevc-ral combinations of steady and alternating loads, Because it would be waconservative to extrapolate an equivalent alternating gtrcs from a combined steady and alternating stres for reinforced plastics, a family of curves similar to those shown in Fig. 5-32 will be required to evaluate fatigue damage. 5-7.2.3 SbvcWW Members Following the establishment of material endurance limits from coupon test data, those parts that carry primary and secondary oscillatory loads should be tested. Components such as the blade spar, whicn may be one continuous member or a built-up section, should be tested thoroughly prior to their incorporation into the complete assembly. Testing of such parts and subassemblies will provide test data valuable for further adjustment and refinement of the streas diagrams obtained from coupon tests. Additionally, tcting oC critical structural members such as the spar reduces the cost of testinl full-size blades or blade sactions. The discussion that follows supplements she test requirements delineated in Chapter 7, AMCP 706-203. In many blade configurations, a full-length mereber such as the spLa. lends itself to electromechanical vibratory testing or other simple loading methods involving minimum fixtures. The extreme fiber stress due to flap bending often is experienced directly by the spar, while - because of a location close to the neutral axis in the chordwise bending plane - the effect of loading in this plane may not be significant. To obtain usable data. the part or subassembly must be instrumented and calibyated to known load conditions prior to conducting the fatigue test. A minimum of three specimens of each signifimant structural member should be tested to ascertain thet fatigue strength ;n the manufactured condition. It is extrermely important that the processing of them "specimensbe identical to that of the final production unit. The data gcaerated from part or subassembly tests will compare wi0i,, coupon data discussed previously, The shape of the S-N curve for most metals is shown in MIL-HDSK-S. A rational method of curve fitting such as is described in Chapter 9, MIL-HDBK-5, steaI be used when no reference curves are available, The diata obtained by using an electromechanical test machine would be simple alternating strem (zero steady strut). These data can be used to refine the Goodman diagrams in a manner similar to that 5-56
.
shown it Rcf. 45. Additional data from tsts of other parts of ilih same material may be used to define further thI allowable fatigue envelope. In many instances thde retention holes of the root blade fittiag receive special processing. Tert data to substantiate the endurance limit of the root retention fitting can be obtained by testing the individual fittinS rather than the entire blade or root section. The effect of bearingizing, (a special rolling treating of the bearing surface of a hole) shot peening, or other such treatment to improve the fatigue life should be evaluated at this time. Component tests of specimens selected as beyond normal tolerance also can be used to provide data to assist ir. the establishment of limits of allowable defects and of overhaul and repair criteria. To the maximum extent possible, tect loads shadl simulate the condition(s) experienced in flight test and be considered in the analysis. However, where well-.!efined stress diagrams exist, a combination of steady and alternating loads that may be converted to an equivalent alternating stress condition should be selected. Testing of the extreme af section member, whether or not it includes a trailing edge strip or other reiniorccmint at wec air icrminua oi the skin. ikucwalc wial piovide valuable data for service life prediction. Because this m,.mber experiences the maximum fiber stress in the chordwise bending plane, tension-tension fatigue loading will provide acceptable data. Due to the relatively sharp contour presented by the trailing edge, failure may be precipitated tt a relatively low stress level by a small nick or scratch. Although the principal stress is due to bending, the critical stress may Wx simulated as'a tensile stress due to the emall gradient. !f practicable, it is advantageous for the test loads to duplicate the predicted stress combination in the trailing edge. Additional parts or subassemblies peculiar to a specific design may warrant special endurance limit testing Among such parts are the tip and/or inertia weight attachment fittin3s. These parts may be subjected to high-aniplitude, low-cycle fatigue resulting from the start/stop centrifvgal force and the attendant secondary mmnent and/or shear loads. The endurance limit for the fittings ann attachments may be confirmed by duplicating the load coaidions cxperienced in service. Other components and/or subassemblies should be tested whenever the construction of the blade does not permit accurate or reliable analysis. 5-7.2.4 Determlaane" of Fatigue Life Endurance limit testing shall provide the fatigue data necessary ib permit the deteinirnation of a ser-
AW~ 7036-2M2 "vicelife. Service life dtermination hall consider, as a minimum, the flight mmneuver end loading conditiors of a realistic misklon profile and the resulting frequency of occurrence of damiagio (ne par. 4-1i. AMCP 706-201). M4
s cycles streg
PROPELLERS
""6-8.1 GENERAL The essential elements of propeller design are described in the pariraphs that follow. Included are a discussion of propeller dynamic behavior and haw it is handled in design; information on the detail dcign of hubs, actuators, controls, and blades; and a description of how test data are used to verify that the propeller has a satiufactory fatigue life. In many respects, the propeller design process is much the same as the design of a helicopter rotor. However, bccause of differences in the technology and therefore in various details of the process, this discussion for the most part is independent of the description in prior paragraphs of the rotor design process. Also. the design requirements specific to propeiers g-eneraiiy arc bcyond the scope of i~s handbook. Therefore, the paragraphs that follow aie 01.ly descriptive of the process and are povided for assistance in thl integration of propellers into the design of compound helicopters, Almost all propeller techiology has dtveloped from design work and experience with conventional aircraft applications. However, the information preseated here is applicable to propelsers for helicopters as well. Where appropriate, there are special comments relative to helicopter applications. Propellers of metal or composite material, with hydraulic means for controlling blade angle, are emphasized. Information on other kinds of propellers, such as fixed-pitch wooden versions or those with electrical blade angle actuation, may be found in ANC-9. The preliminary design procedure for choosing a propeller is described in considerable detail in par. 33, AMCP 706-201. The generalized performance and weight methods given therein allow an examination of all pertinent variables so tLat the best configuration can be selected. The best configuration is usually a compromise that depends upon the relat.ve importance of cruise performance, takeoff thrust, and other characteristics. This systematic method of propeller selection has proven succesful for fixed-wing aircraft and can be expected to provide the basir for the proper choice of propellers for helicopters. "Besides the fundamental performance parameters, noise frequently plays a major role in tht selection of a propelier configuration. If noise is an important design criterion, some further compromise may have
to be made in both performance and weight as quiet propellers generialy require low disk loading and tip speed. During preliminary design, the propeller diameter, number of bladm% activity factor, integrated design "liftcoefficient, and rotational speed will be aelecaed. The planform and twist distribution also will be selected, and the airfoil type and camber distribution defined, The significant performance parameters and the aerodynamic loads under important oprA dine thes rgncocnt terforman earmte and conditiwns then are computed for use in the mow chanical design of the propeller. 5-842 PROPELLER SYSTEM DYNAMICS 5-&2.1 Vibratory Loads The structural design of a propeller is determined primarily by its aerodynamic configuration requirements, and by the structural capacity required to handie the -'rodynamic loads. Although the centrifugal and steady aecodynamic loads must be taken into account, usually it is the vibratory loads that dominate the structural design. Basic vibratory loads originate from several sources, primarily thr following: I. Aefudy-atini 2. Engine (These excitations, which generally are significance with turbine engines and are not disa cussed her. The subject is treated briefly in Ref. 49.) 3. Gyroscopic and inertial 4. Stall flutter. Vibratory aerodynamic blade lords are a result of the propeller operating in a nonuniform flow field, which cacses the aerodynamic lift on each blade section to vary as the blade rotates. For conventional aircraft, the nonuniform flow field as primarily an angular inflow into the propeller disk resulting from the attitude of the aircraft, which varies some with of flight the speed and gross weight. For helicopters, factors that can cause propeller flow aberrations in direction, velocity, and density are listed in Fig. 5-33. For normal flight ope-atin$ conditions, the nonuniform propeller flow field is steady, and the variation in aerodynamic blade forces as the blade rotates is periodic. The blade forces at each azimuthal and radial position may be calculated by standard aerodynamic techniques such as are used for propeller performance computation. The harmonic componeqts of the loading may be evaluated by Fourier analysis of the periodic lotding. Thee harmonics are the P-order atrodynamic excitations - I P, 2P, 3P, etc. - where P is the propeller rotational frequency. Although all of these excitations cause blade stres, the strongest and mc ;t important is that due to IIP (see Fig. 3-.'4) provided the dynamic design of the propeller system is handled properly. 5-57
4ing
MAIN ROTOR
PROPELLER
GINE JET EXHAUST
...
FLOW FIELD AT PROPELLER DETERMiINED BY 1. tVSS.AGE 5. JET EXHAUST 2. M~AIN WING 6. MAIN ROTOR SL.IPSTREAM 3. H01iZONTAL STABILIZER 7. ANTITORIQUE ROTORt SLIPSTREAM 4. VEIITICAL STABILIZER 8. MAIN ROTOR PYLON
Figure 5%33. Propeller Flow Field for Compound Helicopter*
1.
1.0
V,
FLIGHT VELOCITY
I)A - ROTlATIONAL VELOCITY BLADEANGLE OF ATTACK
VTOL AIRCRAFT 0
I
7 T
0
T,
FORCE jOTOF.PLANE IN.PL6J.IE FORCE I'.AiLJT />
/4
f~
RETREATVnING BLADE
~0.6
ADVANCAING BLADE
U D,/ IP SDE F
S0.2
I-
CONVENTIONAL AIRCRAFT
1 2 3
4
E WA
OORV
5
6
Flpure 5-34. Comparime of P-rt Exitations
Figure 5-35. Propeller IP Loads from Nonaaxal Inflow portional to A V1. The dimentionlcas excitation factot EF is derined as
The development of IP-blade loads from angular
F -A
inflow is depicted in Fig. 3-35, where blade #1 isin the retteating position relative to the inflow with reduced an&. of attack and relative velocity and blade 03 is in 0 a advancing position, with inreased angl and velocity. This figure shows that although the resulting loads on the blads and propeller shaft vary at a frequency of IP. the moment and aide force loads orn the airlk~me are always in the same direction. These airframe loads are steady if there are three or more blades. The variation ia lift experienced by a blade section rotating around a propeller centerline inclined to the airflow ii proportional t3 2%e product of the inflow anagleA and the square of tke aircraft indk&aWa airspeed V1, i.e., IP-blade excit&,tion. ic pro-
EF
A
i
48~)
whr Awhe
.
d'Iess,
(5-13)
oelrrelo nle c rorindicowe airglee, deg Ar lentv- xpeso indicated sdt therweodkt AfIn altrodnative excression ise to. idcTeherelationhi oftee P th tweodai excitatsions is A.Terltoai btentetoepesosi Aq= 409 EF (-4 Aj
-
where q
-dynamic
pressure, Ib/ft'
AMCP 705-202
I
WING ANGLE OF OF ATTACK a
" ./V
TILT A/ t A
-- ER AXIa INFLOW ANGLE A
PROPORTIONAL
-%LIFT
2
FLIGHT VELOCITY V, EF-
4
cI.
~2
%.
•
I
W
(3)2
(a--At)
At•.
7.
S32
\
S:
WITH FLAPS
.Y,•-.:•
II
--. 4
',<6
INDICATED AIR.EED V1 , kt
Figrere 5-36.
I .j
IP' Exeltatioa Diagram for Typical STOL. Aircraft in a fixed-wing aircraft, the effective angular inflow into the propeller is a function not only of the aircraft attitude, but also of the wash efTects of the
•• -LVLwings,
o----'-I
SG-L£VRthese
u.0.•
" S-1
.--
C-)
2...
RPLE _ __ __ _ _ _
0
.2dtrie '".
_ _
determined by the gross weight and the wing area, and the slope of the EF line is a function of the tilt of s in dicated , w ing flap s sh ift th e zero-lift lin e an d
fixed-wing STOL aircraft. Th'. ordinate intercept is
_,,,,,,.._
-2_
5W
_.__
__
100
_ _
_ _
_ _
150
_ _
__
bluadeof-h
3
•a
•
the ,•ows ary aaroimation of
arrf e~y ota h vasitonothesbuareofth citation EF can Lie depicted as shown in Fig. 5.36 for
4_
•
fuselage, nacelles, stores, jets. etc. In general,
_(A
100
iNDICATED AIRSPEED V.. kt
therefore the slope of th EF line.) It is customary in factor aircrat deign to consider the I P excitation
ra
._.-'advantageous IP Excltatlos Diagram for HelTcypter Figure 5-37. Push~er Prpeller .. With 2
ytegos egtadtewnWith an •,hen the nacelle alignment is being chosen. higha speed. atFor excitation tilt, the at low factor than no hiher speed may be nacelle may EFdiagram the propeller on a helicopter, itsher p 5-59 ii a
take a form such as is shown in Fig. 5-37. In this case, it would be possible to reduce the IP excitation by tilting the propeller axis so as to obtain virtually no angular flow into the propeller over the entire operating range for a given load factor level. The ordinate intercept is zero because the propeller is not affected by the wing, a; d the main rotor is the lifting "1evica at low speeds. Although Figs. 5-36 and 5-37 - which consider only the I P excitation caused by angular inflow in the pitch direction - indicate speeds at which the IP cxcitations are ,ero, this, in fact, seldom occurs because of the presence of yaw washes in addition to the pitch washes considered previously. In addition, if the mounting of the propeller ib flexible, varation in the nacelle alignment must be included in the propeller load analysis as an aeroelastic effec-t. Once the aerodynamic envirunment at the propellcr plane has been dcfined, the aerodynamic blade loads can be calculated for various azimuthal and radial positions as indicated. However, because the blade deflects somewhat in the pireence of these loads and thereby changes its angle of attack, the actunaakand= asc snighny udii'ircni irom those for a rigid
excitation. The three basic modes for a four-bladed propeller - whirl, symmetrical, and reactionlea are illustrated in Fig. 5-39, which also shows how the engine can participate in the system response. For propellers with three or more blades, respanse to the I P aerodynamic excitation does not involve the nncelle or the aircraft, because, the resulting loads on the aircraft are steady. Hence, a conventional forced blade response (assuming a fixed hub) can be used (Refs. 50 and 51). Such a program must include the effects of the blade tortional dynamics and blade retention stiffness, and dctermines not only the IP. blade loads and stresses, but also the resultant steady loads on the nacelle-aircraft structure. In order to avoid I P-magnification, the blades should have a hiAh first mode frequency and be torsionally stiff. For four-bladed propellem, the 2P-, 6P-, etc., nerodynamic loads excite blade modes that are reactionless with mespect to the aircraft, as shown in Fig. 5-39. Thus, the dynamic. characteristics and response of the propeller blades to these excitations do not involve the nacelle.aircraft system and can be analyzed using the same analysis as used for the I P-cxcitetions. ow dapig associatedO ft However, because of the
blade. The computation of the actual loads must take into account the derivative of blade load with blade
with these reactionless modes, it is important to place their critical speeds, particularly the 2P. outsidt of
angle changes. RANGE
5-8.2.2 Critical Speeds &adReqaose
The response of the propeller blades to the vibratoyaerodynamic paarp6sdtriePyth excitation loads describedtutrl~~1~~/ in the
1
160
and dyn mic characteristics of the propeller syst,.m. The response, in turn, determines the stresses in the blades and the loads and stresses in the barrel. propeller shaft, and the aircraft itself. The dynamic characteristics of the propeller are described best by
/% 120 /
• 100
which the frequency of the aerodynamic excitation
Z
coincides with a natural propeller blade frequency. The relationship of P-order excitation and propeller blade frequency commonly is shown in a critical
,,6
--
" MODE //
U2ND
40
,',',
'
.
iS....
20 0
5-60
-
f
S0 -13;
stronger, critical spees do not fall within the
depend upon the number of blades and the mode of vibration associated with the aerodynamic order of
.
/,/_j
>.
operating speed range of the propelle,. The operating range in the typical diagram of Fig. 5-38 may be seen
to be free of critical speeds up to 8P. The dynamic characteristics of the propeller system
6PI
7,P/
DE
140
defining its critical speeds for the various aerodynamic excitation orders i.e., the rotational speeds at
speed diagram st -h ds that of Fig. 5-38. A'%propellei critical speeds, there may be high dynamic magnification of the aerodynamic loads. Therefore, the rropeller system should be designed so that the lower,
1 1 8P
I I•C~rAION ExciTATI0N ORDEq----, ORDER-v' 8
preceding paragraph is determined by the structural
0 Ftgure S-3.
1000 500 PROPELLER SPEED, rpm
1500
Propeller Crical Spea]Diagram
AMý
1
3@
tho operatin speed raWg with at lemast 10% m rgln. leamt, the values of crtia speed charje with blade anaje due to tht blade twist and centrifuagal effects, blade ange should be conaidered In the evaltuation of reaptionleua mode critical speed relative to the operating range. Also. the effective tetnition stiffness differs for the tiree kinds of propeller modes snhown ;a Fig. 5-39 because of structural coupliPS within the hub ThIis effect must be Included In dynamic and response calculations for the blade. Iii general. the retcntion stik"Nz. h~ lowest for the me actionless modes and h4iest for the symmetrical modes. Propeller aerodynamic excitaiions with.a frtquefl cy order of one greater or one less than integer multiplas of the number of propeller blades combine at the propeller hub to produce bacl~ward oT forward whirl modes of the propeller, respectively. Because Of this whirling action and the rotation of the propeller, these aerodynamic excitations appear on the gearbox-aircradft system as iotating shear and imomfenlt loads at frequencies corresponiding to mutltiplesof the number of blixies. For example, in r' threeblinded propele, ecttooa rqece fZ n ?a w~ elt by laic ovgearo as a or-wnnra.&cu J This interaction of the propeller dynamic systain with the aircraft sy stem must be taken into account in calculating the propeller blade whirl mode critical ONLY OUT-Of-PLAN4E MOTIONS SHOWN IN THE LOWEST EILADE
speeds. This con be done by a complete coupled Amalysis of a rcta&lng fleible propeller atta"he to a stationary aircraft dynamic system. It also cm~ be calculated by first determining the variation with froquency of aircraft system whirl Lapedanoe, e~g., angular and radial deflection oi the propelier shaft for unit shear and moment whbi loeds, and then Including the aircraft impedance in the propeller critical speed analysis. Aerodynamic excitations at frequenck. that are multiples of the number of bladeas excite the piropelIcr In a symmetrical mode, producing vibratory fore and-aft and torque kowd at the uame freluency on the gSa box-icrf sytm Jutaswth thewhr modes, dynamic characieristics Or the aircraft and transmission systam muast be included when smymetrical mode propeller critical sPeeds arn computed. Again, this can be done with a coupled analysis. or by the impedance technique discussed previously. Because the torsional impedance of a transmission system usually is low, symmictrical blade modes that are primarily inplane (putting vibratory torques an tie shaft) will have conside-rably higher critical speeds than would be calculated for a fixed hub. of theJ,cmitriftinl sUMfbaaun cffcct iL. twist of the blade, propeller critical speeds will vary with blade sanle, This effect must be considered in placing the critical spewds pioperly. in general, it is customary to place the lower order wvhirl and symmetrical ctitical spends at least 5%out of tbc nomiral operating range. Less margi4 is needed for thmes critical speeds than for the reactionluss modes because of the much greater damping supplied by structural interaction with the aircraft systemi.
'r- T zOnca SE~NDING MODE.
~
the dynamic characteristics of the propeller
system hove been determined. the maganitudap of Owe response to the various propeller aerodynamic excican be determined. For modes that arc being I'\tations WHIRL excited well below their critical speeds, a real-variable resonse analysis may be used (ANC-9). This always is possible for IlP-acrodynawmic excitation and / sometimes for 2P. System response of the higher order aerodynamic excitations may be determined by an energy method, SYMMETRI CAL using the calculated normal modes of the propelkr and assuming the structural and aerodynamic damping from experience. Another method is to use response analysis, such as given in Refs. 50 and 51, with complex variablew so as to include structural andi L __-Daerodynamic damping. Thec foamer (energy) method uses the normal moJes and natural frequencies ob. REACTIOMLESS from dynamic- analysis of the propeller systan, I'tained and letermines 6ce rsponse of the blade to n particular aerodynamic excitation order by equating the ft"roplle Vatworne Miode 5) Figue -.
-
-
--
-
000011
AMCP 706.202 energy dissipated through damping with the energy introduced by the excitation. Experience shows that the effective overall damping, aerodynamic plus structures, varies with the type of vibration mode, being about 0.02 to 0.04 of critical for reactionless modes, and about 0.04 to 0.06 for whirl and symmetrical modes. From the response of the blades to the various aerodynamic excitations, one can determine the blade stresses, retentioq and s'iaft loads, and, finally, the ',oads applied to the gearbox and aircraft. Certain excitation orders put vibratooy torque, but not vibratory bending moment, on the gearbox; others do the opposite. Also, 2P-excitation on a four-way (fourblade) propeller puts no load at all on the gearbox, as this is a reaction!ess mode, i.e., all the loads are reacted within the hub.
is added vectorially at right angles to the normally considered pitch inflow. Hence, the total 1P-inflow angle is affected. Likewise, the dynamic pressure is changed by the cross-flow component, but for the same gust velocity the wing lift is affected less by a lateral gust than by a longitudinal or vertical gust. Vertical gusts have direct effects upon the pitch component of the inflow angle, the dynamic pressure, and the wing lift. Each of these factors influences the flow field and, consequently, the excitations and loads. As in :he case of a lateral gust, the verticai component is added vectorially to the forward airspeed. This changes the magnitude and direction of the velocity inflow. The current method for determining propeller vibratory loads during gusts uses a quasi-steady-state analysis to evaluate the flow field and aerodynamic
A propeller must have the structural capacity to withstand the combined loading from its response to all of the aerodynamic excitation orders superimposed.
excitations. Although the propeller speed and blade angle may change, depending upon the rise time of the gust, it is expedient and conservative to assume a step change in the inflow to the propeller. In other
S4.2.3 Gaw an Mamyer Gusth and m.ineuvers can heve significant cffects upon propeller vibratory loads. The more obvious
words, the propeller is assumed to be placed suddenly in a different aerodynamic environment without any
change in blade angle or propeller rotational speed, __ .L . ,,- ,.. .,. . p. . .. ,.,i,,.d in the manner discussed in the preceding paragraphs.
effects are caused by changes in the aerodynamic flow
Propeller vibratory loads incurred during maneu.
fields and the consequent excitations to whic'ý the blades are subjected. Secondary effects are the result of gyroscopic motion and inertia forces. Because the
vers are determined by using essentially the same procedures as for gusts, with the exception that the maneures snayr genral are limited toat ose in-
basic frequency of the vibratory loads is the propel-
volving vertical load factors.
is parallel to the flight path of the aircraft, and, there-
maneuvers in~volvintg vcrtical load factors are calcu-
fore. subjects the propeller and airframe to a change
lated using the procedures given in the preceding paragraphs and considering that the effective gross
Aircraft design speciications (MIL-A-8860 series) fer rotational speed, many stress cycles can be acdo not include the time duration of each maneuver cumulated on the propeller during a gust or maneuver. This is in contrast to nonrotating airframe cornnor a breakdown of the maneuvers as functions of ponents, which are subjected to only one major load airspeed. The maneuver spectrum (see par, 4-11I, cycl during a gust Or maneuver. AMCP 706,201) must be available to the propeller Sus' ca co .,e t i -isav .. ........... ,-- n lg-n .... ....... lateral v. Th.....-........... designer so that he can a-.,--- "tat n tcutuc ,,,c u, directions: longitudinal, lateral, and vertical. The the propeller compone,,Lb ,iis be satisfactory. longitudinal, tore-and-aft, or component essentially In the design analysis, blade vibratory loads for in dynamic pressure. The steady torque and thrust on *the blades change with a suddenness that depends upon the rise time of the gust. These changes in load can be relatively high for the large propellers used in
weight of the vehicle is its actual gross weight multiplied by the vertical load factor. Maneuvers influence blade vibratory loads not
V/STOL aircraft because the blades of these propellers are operated at relatively low angles of attack. The change in dynamic pressure also has a direct effect on the IP excitation factor, and the I P-stresses are affected accordingly. A longitudinal gust changes the lift on the aircraft, thus imparting vertical accelerations &ndchanging the wing circulation, which, in
only by changing the aciodvnamic flow fields but also by the resulting effects cfgyroscopic motion and inertia forces In a pullout or pushover maneuver, the angular velocity of precession fl, is equal to
turn, has an effect on the flow field.
where
(n,
l)g V
-
rad/sec
The lateral component of a gust can be treated as a
nz
- load factor, dimensionless
change in yaw inflow to the propeller. The yaw inflow
V
- flight speed, fps
5-62
(S
'
AMCP 705-202 The resulting blade loads can be calculated using a procedure like that used for calculating the response due to I P-aerodynamic excitation. For this analysir, the load is a function of the mass distribution of the blade and is applied perpendicular to the plane of the propeller (out-of-plane). Like IP aerodynamic excitation, gyroscopic motion induces a IP-moment on the propeller shaft, which, for blades having three or more blades, exerts a steady bending moment M on the aircraft as expressed in M - lp., ,fA-lb
(5-16)
where
4,
propeller mass, mass moment of inertia, slug-ft1 W propeller speed, rad/sec Inertia loads result from the vertical load factor applied to the propeller. As in the gyroscopic analysis, the load is a function of the mass distribution of the blad,, but in this case it is applied inplane. The I P shear force F on the shaft is simply -
F = nzWp, Ib
(5-17)
where Wp - weight of the propeller biades and hub, lb There may be other special occasions where loads due to maneuvers should be considered. For instance, a tail propeller of a helicopter may be subjected to large precession rates in yaw while hovering.
J \
54.2.4 Stall Flutter Propeller blades must be designed not only to handIe the applied aerodynamic excitation loads and to have the appropriate dynamic characteristics, as discussed in the pieceding paragraph, but they also must be designed to be free of flutter. Classical bendingtorsion flutter is not of concern because of the large separation between the fundamental bending and torsional frequencies of propeller blades (Ref. 52). However, stall flutter is a major concern because of its potentially destructive torsional vibration. There are two apparent causes of high torsionalblade vibration: aerodynamic hysteresis and Karman vortices (Ref. 53). Aerodynamic hysteresis can cause divergent, self-excited torsional vibration and is, therefore, true flutter. The Karman vortex excitation, however, is not true flutter but a forced excitation. It nevertheles is similar to hysteresis stall flutter and can caue large amplitudes of structural response and possible failure. Torsional dynamic divergence due to stall flutter "Occurs because of the phase lag in the aerodynamic
circulation variation with airfoil torsional motion. Beca.,ac the vortex formation must travel to infinity before full circulation develops, the airfoil angular motion tends to lead the aerodynamic change in moment about the elastic axis. When the airfoil motion and phase lag combine appropriately, aerodynamic energy is fed into the structural system qnd sclfexcited divergent torsional blade oscillation occurs at the fundamental torsional frequency of the blade. Although methods have been developed for analytically predicting stall flutter (Ref. 54), expcriencc shows that a general understanding of stall flutter and empirical relationships usually is sufficient to evaluate whether a given blade design will be subject to this phenomenon. Tests and analyses have shown that stall flutter is dependent primarily upon three factors: the reduced frequency, the blade angle. and the airfoil Mach number (Ref. 55). The effects of Mach number can be combined with the reduced frequency to give the stall flutter parameter SFP. b SFP , ,b, SF, d'less (5-18) aM'w/1 - W where , naturai torsionai frequency,. ad/sec M - local Mach number, dimensionless b, - blade semichord, ft a = speed of sound, fps When full-scale and model blade stall flutter test results for many propellers under static conditions are combined in a plot of SFP versus blade angle, points indicating the onset of flutter form a gener.l trend, as shown in Fig. 5-40. 1 he envelope of these flutter points may be used as a design basis. Although Ref. 55 shows that a blade whose SFP is greater than 1.0 will riot flutter regardless of blade angle or power. Fig. 5-40 shows that, for !ow blade angles, a blade may have an SFP of less than I .0 without being susceptible to stall flutter. Because the design line in Fig. 5-40 is drawn without regard to such secondary effects as camber, thickn'ess, planform, sweep, and center of twist, it is, in general, conservative; i.e., although the SFP of a blade lies under the curve, the blade will not necessarily flutter. In general, increasing the camber and thickness and shifting the center of twist forwurd will increase the blade angle at which flutter occurs. The effects of planform and sweep are more difficult to assess, because the stability of the blade involvts the integrated effects over the entire blade. Thus. although some blade sections are stalled, the blade itself will be stable unlss the integrated energy fed into the blade is greater than the structural damping present. 5-63
I
. ,s. , J
4P., AMCP 706-202 1.1
-
I
j_
S ,0.9
CM, increasing the aitfoil camber is one way of increasing the forward thrust or power at which a blade
-
-
-
-
will be subject to stall flutter. However, in all cases, if the $FP is greater than 1.0, the blade will not be susceptible to true stall flutter regardless of blade angle or loading. The other possible cause of high torsional bltde
0
<
NO
0.8
response, and also bending response, is Karman vortex excitation. When the vortex excitation frequency,
0.7
%-which
is proportional to (aM -VI -AV' ) /h, coincides with the torsional or bending natural frequency of the blade, significant blade response may result. The
(L 0.6
i. S0.5
frequency expression is the reciprocal of the SF*? by the torsional frequency, i.e., (SFP/c,)-1.
-divided
Although the coincidence of the natural blade tor5ional or bending frequencies with the Karman
FLUTTER
I0
0
vortex excitation fregqucncy can result in significant
blade response, particularly at very high blade angler 0.3
1of
S10 15 20 25 30 BLADE ANGLE AT 0.8 RADIUS, deg
Figure 5-40.
35
Stall Flutter Design Chart
attack or blade stall, the response is not divergent.
This type of response can be experienced at high blade angles even by blades whose SFP is greater
SHEAR STRESS
than 1.0. App?:¢ent from this dikrucsinn is thr. d--sirsphility of propelier blades with high torsional frequencies so that their SFP is greater than 1.0. If
tA MEASURE 01?T
this is done, divergent stall flutter is avoided. Also,
STALL FLUTTER)
the higher the SFP, the less likelihood that Karman vortcx excitations will be a problem at very high
__
_
-designing
--
C --
blade loads. In general, solid propeller blade construction, e.g., using aluminum, gives rclativcly low torsional natural frequencies and SFP values below so care must be exercised that these bilades are not operated in the stall flutter zone of Fig. 5-46. Although composite monocoque blade construction
]1.0,
1CM
struction, it frequently gives SF)' value,, below 1.0. In
-
general, propeller blades consisting of a structural
0
2
4
6
8
10
12
ZERO LIFT LINE ANGLE OF ATTACK, deg"
spar and a thin composite airfoil shell have SFP values of inu.h greater than 1.0 and are not susceptible to stall flutter. 5-8.2.5
Propeller Roughness
Airfoil Characteiltldcs mnd Stall Flutter
A propeller can apply loads to the aircraft with resulting vibration or roughness that is unacceptable
A blade also may be evaluated for stall flutter by analysis of the characteristics of the moment cocfficient Cm of the airfoil. As shown in Fig. 5-41, the onset of stall flutter usually occurs when the airfoil is operating near the peak of the C¢ curve with respect to the torsional elastic axis, and not the lift coefficient CL curve. Thus, stall flutter occurs only at high thrusting or blade load conditions - both forward and reverie. Because, for a given blade angle. airfoil
to the aircraft structure or the occupants. These vibratory excitations stem primarily from two sources: excessive wAss and aerodynamic unbolance of the propeller, and an undesirable combination of nonuniform flow field ano propeller dynamic characteristics. The former shakes the aircraft at a frequency order of I P, whereas the latter shakes the aircr'aft at frequencies that are multiples of the number of the blades - e.g., 4P, 8P for a four-bladed propeller.
camber increases the thrust or CL without changing
Through proper design, manufacture, and assembly
Figure 5-41.
5-64
..-.
-......
to the blade being offset from a radial line. Toe mo~ment is transmitted from the W~ade to the retention, to thu bariel arm, to the front an~d rear rings, and into the tailshaft. The inplaf,4e cokiip:nent is reacted by the spianed joint, and the out-of-planc comnponent is miacted ela'ttcally within the barrel. 3. Propeller thrust. Thrust is transmnitted from the blade to the retention, to the bladc arm, to the front 4nd rear rings, and to the tailshaft wheiv it is reacted by the tajilhaft thrust bearing. &-8.3 PROPELLER HUBS, ACTUATORS, AND 4. Propeller torque. Torquec is transmitted from CONTROLS the blade through the retention to the blade arms and The design methods desriked in this paragraph into the tailshaft, where it is ruacted through the drive * relai.c pnimarily to the propellers of one wnanufracsplint. turer. In these propellers, pitch change actuation is 5. IP-acrod)nainic bending momenct. The inpmanc * hydraulic, but many of the design aspects a~pply to almost any crinfiguration. The discussion is limited in component of W-vibratory loading is transmitted from the blade w'.the retention. to the barrel arm, and scope inasmuch as propeller d&sign, development, into the front and rear rings. There, because of the and manufactare usually are per formed under subun~ymnictrical load phasing among the blades, it is contract, or wimder sepa.-ate prime contract, and proreacted elastically within t.he barrel. Thie out-of-plane vided usi Governmzrit-furnished equipment (~GrE). contributes to a combinett bending moReenditscomponent BarelsodBlad ".3.1Proellr wnt on the tzi!sraft, reacted by front crnd rear rvdial ropelerBarel ad BadeRetetlos 5-8..~ The shape of the propl".e1r ba; re ibdetermined by bearkings. of th%: propeller, the detrimental effects of propelicr rou-,hness can be minimized. Tht importance of propellcr balance and higher ordcr excitation to aircraft roughness is dependent upon the propeller mount design andI its integration with the ovevall aircraft dynamic system, which determines the damping and transmisaibili~y of these excitation loads to the airframe.
~ ~biO3 ~
)high -
-.
~pl,
~twppe.~ ~
~
~
power pvopt~lers with solid aluminum blimdc~s; a one-jiiece barrel wil"I a single integral blade retention race. is used for low j)3wci propellers with sofiJ aluminumr blades; a onP.-piece barrel witli nidltipc 'ritegral ra#;-ýs is used fi'7 propellvas whose bizades are made with a steel core and a Fib~rglas shell. The~ iast type g'-nerally kisused Ior he~licopter installation arid is the r dhcui~. ipe hat the barrel tatseiibclyS. t--74. &te areb Thearrl co.!its asemby ik'te brrel bidt: retkiltiun balls, seals, and clampe. I he barr-zl itscif is a one-pi=ce vacuum-nclttd steci forging %Icorpoimating the bcoiring race tor blade retention. The tziiish~aft of ihe biarrel exttends into the gearcws and is di-iven dircctl) through a spliaed joint. 1'1e tailshaft ir supporleOi cnm two radial roil.ci ".~rings - front aijd rear - so that pLropciler moments arc reacted directly into the gearcasu iiousing. Propeller thrust loads are reacted through an angular con~a,:t beýring at the front of the tailshaft.
"4.3. 1.1 Barre Loading The following lords arm conuid_ýred when dcsigning a propeller barrel1. Centrifugal force. Blade ;entrifugal force is transmitted to the retention, barrel arm, aud iront and rear rings, where it is reacted clastically within the barrel. For purposes of analysis, the barrel isconsidered to contain a front and s rear structural ring, 2. Steady bending mioment. Steady bending imoment is due to aerodynamic loading on the blade, or
~
Y Ih-acrodyflamic s~dc Formc. This fivuic Is tri-as-
witted like IPout-of-plane moment end is reacted by the tailshaft radial bearings. n oet ihrodrvbaoylas ~ 7.c athigherorders (2brtor loadset.) ombints ind ociahgerrds(2,3.4
t.)omnen
various patterns depcnd'zmg upon thie load phuming amnong the bl,%&S. Accoldirig to their patterns, these leds may L-e reacted elasticaily within the barrel or
transmitted to the tailshf~ft bearingr. =d :=t'd as a bcndng moment or a fori.-ut,d-aft tur side force. They
as obn navbar oqe als cyombinpin voribrntr toqe.dammetfo gyrc rscopic onafecs moment. ndei.arr momen fParomy ieIPar~y cinafcstebre grsoi nm~ oet 9. F-ropeller effective weight. Tht F-ide force from propeller weight, rmultiplied by the aircraft vertical lodfcrafetthbrrlikIProdcwc side for.:ec. 10, Bearing precssi~t loads. The t&61shaft bearings are presshrt onto the t,..sshtft and itaposc a compressive load locally. It. Axial preioad. Tensile load is imposed on the ta~sbnft by &Aial prcloki.ding of the shaft bearing agsinst a shaft shouldet. 12. Blade t-vi. ling moment. A combination 01 ccntrifugal, frictional, and aercmdynarnic :ffects, prodLczs a twisting moment around the bildc pitch change axis. -This i6 tran~fcrred indirectly to the barre' uL a couple beten the bWade rctentior ant: the pitch change actuainr. 5-65
- ý z '- 'ýýý .j;ý , -_!_I ý
..
lie
543.1.2 Lcadl De~hdwh In designing the barrel for a specific configuration, the various loads described in Mte precedig mahst be defined for the applicuble aircraft 4parag~aph mission profile. Loadings idivolvod in demonstration and qiialification test aiso must be .onsiderect. Applicable specifications hall be defined by the propelle.~procurement spvirtion. From the analysis of propeller design parameters and dynamica, a summary of sigiiificazfl loads is prepared for selected conditions. These loads are the basis for tl~c structural analysis of the barl 541,3.13 Barrel Sreetaal Tests After the barrel has been manufactured, th stresses at critical locations arc measured under various load conditions to assurc structural integrity. The barrel isari-gaged, and m exermntal stes analysis is perfornad with axial loadis applied to simulate contrifugal loads and bending moments, both static and vibratoc.~ An assembly with cylisidericil test bars instead of so tkah retention rpdg rate can be deduced. IW&is scomp"re with thec ciakculated values. The barrel she isfatigue teutest to determine its actal muria of safety under the desWu loads.
'
~.Y
---
1V
-- 1 .t
7
system or from the propeller itself. and transmits a signal to the. actuator to change blade angle ak Iiocessay. The actuator converts this signal into a rnichanica! action to move and mai.ntain the. blade angle. 543.31 CeftWlo Comigufradosi Several arrangements are used for control coniponenta. In one common conr4uration, the control assemblj ir nonrotating and mounted near the propellcr on the gearcase or surrounding the. barrel talshaft. The outpat of the control is a hydraulic flow that istriansfarred to the pitch change actuator in the barrel assembly through transfer bearings. The rate Of flow controls the rate of blade angle motion. In another configuration, th-. staL~onary portions of the control produce a mechanical signal directly rclated to the desired blade angle. This signal is transferred to the barrel assembly through a mechanical bearing or a differential gear train, and the rotating astarnbly contains the hydraulic valves and pumps required to drive tha, pitch change *,uator. C~uhtcu..qpeW Go
frqoy
5L.2.1
4.4L3.2 Fr"rAs Aedsosamned (entralmi Propellr Fitcti chrnje actuators and control ~seenas adjust anad mai stain blade angle according to one of several montol wamod as required by aircraft and engine opeaouing condition. One control niode commionly used is constaut-sp~od governing, in wh*a adacWe propdwir rotational spaca is hed wmn~an by a governor that raime or lower blade antge iw resioosc. to changes in forward speed or applied powa. Thi is the contral mode used almoet uastiuslly int flot "peration of comisational air4caft. wk. it alowt fthpilot to scec the most efflesent cembinatiiim of propeller wnd cagne operuting conditions. The other conanon control made is buts control. in %ssvb ýhe contiol aut a sclei4ed blade angle 0l. ina rapoane to direct pilot control (TY to fte output of a coordiuated engine or aircra'I fligt contad systemz. ket control .inis used for propeler reversal during land ingn kaw-thrmst ground operation. ard for propellers uWd as primary aircraft ligh controls, such as in VTOL aircraft. Many propellers have a combination corkrol system that uses either constant-speed goverwning or beta control as required. 1Ue pich contro.l sygatr has two basic connpowtnctLts. 9 cool aol and tki actuator. 19hw control rsceivcs ujuaals frome the Pilot ar another control
The input to a constant-speed governor is a si~agnl( calling for a d"sred propeller speed. This signial may conic dirwcty feorn tha~pilot or from another control system, such as a synchronimc or a coordinated Power inanaemoznt control. The deie sptd 'a compared with the actual speed by a device such as a speed-set spring balanced against a set of rotating flyweights driven by the propelle or enginec. If aai offspeed condition occurs, thet device puts out An error signal, commonly in the form of a pilot valve dispiacemnent. The valve dusplacement. in turn, meters oil to the pitch change mctuxtor, raising or lowering the blade angle to sdow down or speed1 up the proPellet to correct the off-speed. In forward-thrust opriatiota. blade twiaing momeat alwdys is toward low pitch, and the actuator loads always are in one direction, The governor, then, nwa to mete: Iuigh-zprmsure oil only to onc side of the actuator piston to raise the blade angle; kwteizg oil drain back to the stump permits the twisting- monient to lowtw tbz blade angle. A governor widl only this function is called qt sinae-atng governor. A donbie-aciing goversior can direct high-prsrsure oil to either side of the &atuatorpisto. Ths cqwzbalty is required if the bModes are to be controlle in revaeri thrnais operation or to be uWeathered. In many aonstant-speed governors the pio valvt positioned by the upeced-aonsing dlevice m~etem sctilator 041 flow diietiy. For some designas. bovrmi,~ lawg oil flova arm required and ~an of a sample pilot
i
.
.
.*e
-
AMGP 706-202
N valve would cause large hydraulic forces. Speed acnsing-accuracy will be affected adversely unless correspondingly large speed-set spring and flyweight forces are used. To avoid the weigbt and size of such a design, a servo-type governor may be used. In this design, the speed-sensing device positions a small pilot valve that controls only the flow to a servo piston, which, in turn, positions a servo valve that meters the main oil flow. Because the pilot valve is isolated from strong hydraulic forces, i; and the speed-sensor can be made light and sensitive. 54.3.2.1.2 Beta Contr" With beta control, the input to the propeller control is a signal calling for a desired blade angle. In recent configurations, the signal is transferred mechanically to the rotating components, where it positions a distributor valve spool. The deeve of this valve is positioned mechanically a a function of blade sngle and, if there is a disparity between desired and actual blade angle, the distributor valve directs high-pressure oil to the appupnate side of the pitch change actuator. When the blades change pitch, the
other propeller designs incorporate an independent hydraulic system. The complete hydraulic system contains the following basic componenti: a sump, a pump, a filter with a bypass v&Ive, a relief valve, the control valve, and the pitch change actuator. Schematically, the actuator is a linear hydraulic piston with a meohanical device to convert the linear piston motion to rotary motion of the blade. The maximum operating pressure (relief valve setting) may be 1200 psi for a simple systcm or up to 3000 psi for a system where weight is critical. A schematic diagram of the complete control and hydraulic system for a double-acting governor is shown in Fig. 5-42. 5-8.3.2.3
Auxiliary Fuidoam
control, some propellers provide various auxiliary functions. Two of these - feathering and pitch lock are safety items. Propellers arc. feathered by turning their blades to a 90-deg blade angle - i.e., edgewise to the relative wind
valve slekxc moves as well, until the desired b!ade
- in order to bring a disabled propeller oi-
.iijit a
angle is reached and the valve is closed.
liop nuiin .ui
l
The pitch lock acts to prevent further decrease in
54.3.21 Hydraulic System Some propeller control systems usc engine or gearcase lubricating oil, boosted in pressure by an extra pump, or oil from the aircraft hydraulic system. To minimize contamination and improve reliability,
pitch travel when normal blade angle control is lost or a preset maximum rotational speed is exceeded. Without this feature, ,a loss of hydraulic pressure could allow the blades to drop to low pitch in flight, causing a dangerous windmilling ovcrspeed.
DESI•ED RPM SIGNAL
DECREASE
BLADE
PITCH
MOMENTSPEED
-
/ -
ISP
RPM
SET SRING
YOKE & TRNINfDOUBLE-ACTINGI ROTATING
-~
_
FLYWEIGHTS TRANSFER BEARING
L___
DECREASEI
PITCH
0.
GOVERNOR
I IDECREASi_
E
PITCH
FILTER AND
.BYPASS
VALVE
II
NUr
ACTUATOR
_A
PTHCHANGEWAEW ACT UATOR
I
SR6 I__
____E___
Fhgue 5-42. Prpodeler Control System Sebematk 5-67
S-43.34 Coutrol Porlbasaevi UP?anticipatory sigal Ebr 4*otrol thut helpa to minim- o'vershoots and provides good systm u.ýabiiity. A propeller control, in combination with thc ongino conirol. must set and mnaintuia the opeiirstihg 54L.3"2 CoaWrl RdkM t~y conditions directed by the pilot. It must relipenduscotl fntiscacus Beas otowafntoscncuearious quickly and accurateiy to pilot commalids and be unaffected by undesita~ze disturbances. Contrul Wtrot'ble, reliability is a vital part of propeller coutrol design. There are two diferent philosophies iuavoived forinsancc is summarized in three parainew~s: aicinl design for rtfiability: aafe-life and fail-safe. curacy, stability, and transient response. In order to T'at safe-life theory requires that thec probability of defiae thex.e dharacteristics it is n oesary to study the behavior o. the "ntire propulsion system over its a catastrophic malfunctioa be uacecidingly rmiote. The dcaign m.uv, have n mufficient margin ofsafety for complete operating tange. using vdr our analyticni all operating conditions, both normal and abnormal. techniques. PrOOf that A system enJOYS this level of reliability The ovvrall transient performance of an aircraft propulsion system generally is analyzed by using a requires extensive testing and thorough analysis of *sa-vic e~rincflC. nonlinear dynamic simulation of Vie overall system. The second methol fai-safe, requires; that no This involves detailed nonlinear dynamic equations for the engine slid the engine control, as well as for- reasonably probable single mal1function be allowed to cause unaatifactory operation. The design must irlthu propeller control. This simulat~fm rquires input.p clude safety devices Rnd redundant womponents in that define the ambient condition.. surrounding the enaine and propelltr - such as pressure, tempetaOrder to tolerate single failure. Some Pr*pulcrs have lure, and flight speed - and the pilot-initiated inputs dual hydraulic systems, both controls and &c~uattirs, to provide fail-safety. This is the case espoculaly for that specify the desired operating point. This type of propellers used for primary aircraft control as weUla prograin provides estimatts of thrust, propeller *,%--A
cna
_A
A**~.*I A_
~
..
ables as a function of time for various types of pilot command signals or ex~emnal disturbances. Fig. 5-43 shows a simplified block diagram of a turboprop propulsion system and indicates some of the typical input parameters to the engine and pro-
peller controls. Each of the blocks contains a matrix
of nonlinear diflermatial equations that ame used to define the steady-state and transient behavior of that component. no A simpler analysis, providinq a good inaigiht int the basic control roquiý can be made by line-. artisng tho importanm .ystem parameters at various discrete: operating r~iditions and examining the systemn behavior for small diaturbances around these operating points. This pennits the application of classical servomecha-nismn theory to the design or the control system. Fig. 5-44 shows a typ"c linbirized block diaram for constant-speed con~tol of a turuincv4riven pro. peller system. This integral coilowl system wail move the blade angle unail thre speed wrý Sons to=ro thLA assuring that sensed speed is equal to desired speed under steady-statc conditk'..;. The governo shown in Fig, 5-42 is such a conu&.:. the control designer may provide gead ,oaipuasatoa to improve the speed of respons and tie- stability characteristics of the symcner *f kf had charmaktrsi compenstes for the kg tirý vnamrt between blde angle change and speed chang. la efect, it provides
for
r
An importaait part of design for relii~lty is the Failure Mode and Effect Analysis (FMEA). Such an analysis will help to revcal areas noeding better *Afety features. lIt als wiill provide an ;inprover, understanding of possible malfunctions ao tha their (ton-
sequences on aircraft operation can be appraised
joinstly by the airfirame designer and the propieile manufActU-r. 544 ]nPROPLLRmBAE
S-4-41 Wad* C08staby As expiained in par 54.1. the aerodlynamnic size of a propelle is chosen initially with the help of pammetric perforivaloce 1.widme in which the trends of performance and weight are evatluate for various combinations; of chacactaristics. For the paruncuric studics. a number of important but sam.edary &erodynamtic design details arm assumed to he in a -standard" condition, to be evaluated later. For the complete bLde design. the sazandary surodynaw. ; deails - %ce., t'*Aness planorm. twigt, airfoil type, and CAM-ar - are Chowen f&ra rom p"s expetWmm. Major consideratio is given to the anticipiated severity of struictural loading and the type of the specified operatin coeditiosus - aerodynamic performaence and loadig, and the blede structural resporse to the lads, ar e mputed. Wiah thene resut&. dhavong ame mak in the desig ;"kal aid the smalyses a:,- repeaed until the desired leisk of aero-
1
D
CCL
PWRTFIE
GIEENGINE
4.
PROPELLERPELE
>n
INGNLET IMJ e.
IL,
IUI]. ENGINE
PROPELLER CONTROL
___CONTFIDL
rq- S4
UhpM
-s
-~balk Dilwa
_10'
"gm" CA, aN
CH. NGE IN PI
ELLER TOP
WIn
ENGINE
J -'
•.JE W T N, , H
addition, if the frequency of aerodynamic Ic.4d approachti one of the natural frequencics
"variation
a0,
4-12 CHANGE INPR)PELLER TOCE WITH
tADL ANGLE
S - LAPLACE OPERATOR
a
3. Bending stiffness
P
5-44.
+i1 s
IS 0 rS~'
5. Bending frequencics 4, Torsional stiffnles A +-
NA N
Lhedued Prapeilhr Coo"d Mark Dagr•
dynamic performance and blade stress are achieved. For effliiency and .esd these analyse are performed with the hep of a high-spc digital d xrneputtr. Of the secondary details, thicknes and planform arc the most important to propelLT strength and weight. The choice of blade thickness and its diribution alor4 the blade is affected by structur -I capacity requirements and also by consideations of vibration-caitwcal speed locatiom. The aerod),namic remquirement to minnuinze ptofile drag losss puts a conszraiat on blade thickness, and specil pr:•blcs such as tip compressibility also may affect tO design. rh,. nWlaA.o ^f~ir n~nedan arn f,_hncneaffnI hw the choice of blade materil ad construzbion; with sparand-shell Fib•rgla bladeis a tapermd planorm (wider inboard) lnds itself to both minimum weiht and redimced noise. The slcted activity factor usa", limits the amount of taper - the hihe the activit) factor, the kas the taper. For solid aluminum blades, a planform with an elliptical or rond tip oftrn is used for reduced nse. S-S4.) Ebb Cemnbudl Viewed as a stracture., a proper blade isa rotating. cantilever beam subjemtd te two major chumes of loading: inertial forc and aerodynamic forctm. The incrtial fore consist of steady centrifugal loads and vibratory rectiona; the aercdynantic form are
both steady and priodically vrrjing. Steady loads am important to th blade den, but vibratory loading usiadly i the domiinat ianuemnc. Even 5-70
cyclic loading in the blade can result. Much effort is magnification by properly positioning the bending and torsional frequencies. In gencrul, there are six major structural asp(cts that must be considered in designing propeller blades: I. Axial load capacity 2. Bending capacity
00 aN, ;'N'
FIip
of the propeller, considerable magnification of the
spent in blade structural design to hmit this dynamic
JO,
i
without dynamic mrz4-oj -tioia, vibratory loading at one cycle per revolution (I P) can be v wry significant.
Toersionaloffrequency. 6, ot-jective The the structural designer is to cbtain, by judicious use of various materials and configurations, the lightest and best stiructure posible within defimed by aerodynamic r geometricson.trait the
4L4.,.1 Types d bWa Co isuutew The i trn•p • iiuv 6€, the solid b 1 C ti shown in Fil., 545(A). With the propec choict of material, this coostrwsioa has provided an aorxptable balance of wenit and straure for many years ort conve•niall ai.rft. d its advantages ar&simplicity and low cost. Its diaLAvanta#cs are ineflicient usc of material in the cente of the cross section, partiulaty in bending, and the fact that the primary structure is exposed to service-inflicted forign object dampge (FOD). Unfortunately, proprller blades often are subjected to the impact of varkiu; objx"ts, r ~n.;n. frnei taimi snAi d tn sinr mad birds Many of thke impacts arc capabl of inflicting surface damage that proppag as fatigue cracks because of the r~ydic sztmes caused by the untcsady loadings. Recognizing the potential of FOD has been shown repeatedly to be essential to acceptable service efcmanc. With solid Construction, the designer must select materials with low notch sensitivity and low crack propagation rates, and must observe conservative stres limits. The simplest form of holio.;-baide construction is shown in Fig. 5-45(B). This is a fullyatresscd skin of monocoque construction in which the central material of the blade section, which contribute, little to beam stremgth or stiffnes is absent. The simple hollow blade construction has the potential for sub-
stantial weight reductior, but two significant problewa have interfered with mrusing this potential. First. foelvn objc that only gou•e the surgfe of
"
gougo and dent tih hollow sec. can b tio, The dent causes an additional conouttration of viewin the same area as the goup. Therfore, the nniimum wall thickness is determined not by grow structural considwations but rather by the required restano to FOD. Becasis impact velocities are greatet at the tip, thick walls am required there; from the standpoints of oentrifupal load and frequency placiment. this is an undesirable region for added weight. The meod prohlem ef simple hollow blades ocours at the sedn and traLing eds which form the only shear lod path between the thrust and cambe- fhee. Because the faces interamct at a sharp angle, thd.bend radius is small. The resulting geomearic stres concentration often is unsatsfctoty, and additional material must be added at the edes to increase the radius and reduce dte local Utrns. This added material not only increases the weight, but also reduces the blade ntural fmqueame, prticulaly in Ssolid blades
Lorsion. Soine of the problems of simple hollow blade sections an alleviated by a modified monocoque conu m. 5sunwtion., such as the ribbed section shown in 45(CQ. Osmc or more ribs am Widvd to supply13 additioral shear kad paths and to reduce dmtin. T euact proportioning of sheet and rib thicknessma spacing along dhe klnth of the blade is a co•bina
"
•
""'
(A)SOLID ,_:;
(B)MONOCOQUE
lion of analysis and experimetntation. Once perfected, the resulting blae generally is much lighter than a solid version. The primary structure, however, still is exposed to FOD, and cracks caused by such danmag can propagate acros the entire action. One way to solve this remainiqg deficdicy is to build what is in principle a modifed monocoque notion in two pieces. Fig. 5-45(D) illustrates the caoi section of such a blade xistruction. In this approach, the central tubular mmber. or spar. L mad as a tapered-wall, varying-diam er tulb and them is flatmt d and twised to shape TIh oter aerodynamic contour or shell taen is bonded to the sVpr. This contruction aows consmderable design bility. The spar and shell wall thicknesss can be vared indep=en9tly to achieve the required -016 and stiffness distribution, and th. spar position can be varied as well. Additional within the l ftfium~t, suc as choAdwise wall thickess vaniMos important, with the separetion of the aerodynamic shel and structural sqpr, different materials may be considered for each item. With the proper ahev C =hAi=4, 1tbini s-3i6 i.e., the effects of massce inflicted damag can be limited to the shell, elowiun adequate time for detec. ton and repair. A typkial spar-shell blade is shown in Fig. 5-46. Necaust the apur is foamed from a tube, the transitie from the airfol contour to the round c-etieo,
(D) SPAR-SHELL
SF-..
K
tinuous from rettion to up. with no Joints of any
sort and th sh
i bonded to th• pm over t hen-
tre blade lekth. With the lag jouit ar and the of load from the shell to thk mar continuous tani ieh bonded joint isloads ihtly. A filler ma al. oftoa is md in the cavitin of the blade. This type of blade comAtuction has ben very satisfactory, with t nowm wCk dim solid Wlades. b T(C we of holow blades to adtcive thesemlprovmnents, however, introdues a number of special structuralconsrations in addito to the ax major
no batwm
sbhw and airfoil. wher lo
wl-
ing of the spar and 6ill wallk can occur w=m th, eNd. to straigten or dform the• kxmos b=di tudnlbW profile of the walk. 2. Panal vibraio of local area of the spar or shell, especially if they anr Uisupported 3. Shear flow, frhm the inailty of the mqx sbhsm name the Sioad in a hollow blade to ram diey
Wande thicknessm. Instead, dimgg lead must foallo .,, Mae& Cnsm Sedliss ' ': •"Ty*pl 5-4
-
section is natural and convenient. The spar is •wn
I. Secondary stractural action in die transition
-a*
.&
aU4tioni both shell &adqw.s also can be achived.
aep-Als mentionAd previously. These am
(C) MODIFIED MONOCOQUE
•',
the caits th~lm~m/fa dedetbS~m•,lm~t material emed strictural
dIe -7t1
7 .xw -:-
AMCPC blades, especially, deltail blade design must represent an intimiate combination or functional requirement / ~and manufacturing capability. Monocoque bladcs are made by manufacturiag s procedures that fall into two general categories: flattening and twisting tubes ino shape, and jouinmin a I * number of individual pieces. , In the tube-flattening process, a round tube is drawn, extruded. or reduced in sonme manner from the billet so that its wall thickness tapers as required * * and its perimeter is compatible with the desired airfoil sections. The shank region isswaged down to the NICKEL EROSION SHEATHproper size, and then the tube is pressed and twistd SHEATHinto shape in dies. If the blade is to be tapered, the trailing edge os' the pressed tube is trimmed and welded. Some hollow blades arm made from only two pieces, one containing the shank region and one surface of the blade, the other the second surface. The FBEGA%-OEE FIEGA -OEE two pikice are welded together at the leading and
7
-~
~'
"
-V
. -
LOW-DENSITY CUFF METAL SPAR
FIBER0LAt SHLW FIBEGLASSHEL
-~ ____Hololow
LOW-DENSITY FOAM Sw~il lo&The Fkmv54LTyplcii F~weS-4~Typial SM*~Ncording 4. Filler streseing fromi carrying some of the sher ris loading across the blade thickness and fromnuo to cross sectum deformation S. Shear lag effects, in which rqom of the sufav distant foom a rib tend to duck the buiding load 6. Chardwisc deformation and stress from and inertia Woe"i on tOw airfoil, ab well as pnv from nlawIg effcts 7. Stmcutral sabl. .y- fte avoidance of budthg I. Joint struaing especially at the Anboard end of the shel. These possible problems must be considered Carefully in &una, and reouir judicious choic of configuratin., filler, and jinit design. Additional ifformaticim on type. of ptopeller blade contrctininluding both wood and metal blades, may be found in ANC-9. I&Awelded
Mere so than wub mansy structures, the demig of 48, mtrom proPsller~ bla es h onhem aFfecri by the ",oufnasnngpwessusavailable Zo olw 5-72
trailing edges, and abrazed fillet added to strengthena
thr edges. The pieces to be assembled are milled, ýgru..ad, and press-fornied into the proper thickns eort jo...... wa&.0 h~masav bftm "Wd rrom smaller sepmets, with both cliordwise and longitudinal welds. bladies with ribs ate made essentially the sawe as the blades just described. In this casw, one o moreof the pieesttobe weldedor brazd isfmilled t contain the central ribs. manufacture of spa-shalblWades varies acto the material used. These bkdade have been produced both in all-metal conr~guations and in configurations with steel spars and Fiberglas shells. Current development work is demonstrat*n the promisc of new comnposite materialk for the spar or the shiell or both. Metal spans are manufactured by processe siilar Withose ased for flattened-ube monocour. blades. Interal and exernal psening with metal shog or glass beads are used for strengh improvement. The retention bearing raceways Itme are machined on dhe root of the spar. Thene raceways are inegral with the: spr material, and mugt be hardene locally by carefully controlled Rlam or induction methoci. Finally, the: rinished spar pay be plated for corrosion protectioin. Metal sheb arm formed froui polish-ponwd shemis, which ame ROWd arouud the leading edp and seamtat the trailmig edge and the tip. As weth the Fibergla shells nonnally are l6i up on metal mandrels, either by a wet inyup proces or by uwsai prejials ab" Sm igoblinebys" wetam*
1
i
"
.-
IIIf
.--
-
-
'-
A"c 706-20
in .ne piece, folded around the leading edge and open at the trailing edge which isjoined when the spar and s are joined. Others are made in two pieces, with bonded joints at both the leading and trailing edges. TUe pr and shell are joined by brazing, for allmetal blades or by adhesive bonding, for the FiberSlas shells. Sraaing is performed in dies in a brazing funac, with pressure in the spar to hold the joint in intimate contact. Adhuive bonding of Fiberlas Su may t be amcomplished with heat and pressure in a pair of es but the more common practice is touae a die only for the camber faces of the blkd and to apply pressure by vacuum bag or autoclave. The ~asembly of a spar-m ) blade also reqirmes
shot-peering, arm subjected to thorough process control. This includes fraquent inspection of the machines and techniques involved, as well as priodic destructive exa•ination of sample blades. The surface condition of propeller blades should be subjected to careful visual scrutiny. In addition, sparto-shell joint qual-ty is assured by process control and by nondestruaive testing techniques such as x-ray, ultissonic scanning, and tap testing. These nondestructive methodi also assure the quality of the blade filler material.
ssseel a"enjoined. ast-inns ipr taled asat j ofnd s •'then awe filly ais Wsed she tba is pthredinto blade
S-t.4.3 Bid m and Proeller Satate Propeller roughniri due to unbalance is caused primarily by deviations from tolerances in the three mSen propei compone blades, hub assembly, and spinner. The most impoitant souror, of propeller unbalance is the blades, because of their larr. radius
joint is made. after the spar-shell cavities aad cured hels coposte axeiah in O adanc Spasam Sbe presnst ate of be am usually an laid up by
aerdynamic characteristics. and maw and their Four factors aff'ect Ong amount of blade unbalance: mass force, mass moment, aerodynamic forces, and
hand in tape form on mandr". When the composite
aenamic rment.
oliihtwtisit file material. Some f'lkir installationof or or cut from piom am ppecast bonyomb; theseto shape ae installed as balsa prt wood of the honycmb;y wlthes
maiMX is rnaU Ste
. mrmaina G
.. I.• .
W
.
as a die and autodave Prusum or with an o'tsik d and inflatab bladders inside. When tft matrib is the mtea is compucled and diffumon"meta bonded in matched tal din. With advanced ompit,
materials and
To obtain
ealiuic estimates of probable un-
balances for a partcalar propeiatr design, the ©tffM.. of t varion independent dimensional toldrances must be ev-juad stalieically for each of the three major componmts and then combined statistically for the overall propeller. Analytical techniques and
fiber cantationa can be umed in thn various layers of the spr and shll. Namerically controlle tapelayins mkckins an beit useci in curent deveovt manuf'acture nmnt progsams to facilitate the cii of thes new blades.
computer programs hzve been developed for cmtimating and assesasig unbalance for numerous propefler installations with good wss. A discussion of the varoa. aspects of propeller balance is Oive in Ref. 56. AAditional comments appear in ANC-9. In pz'rral, blade balance is achieved in manufac-
The prodoction quality of a pgopele blade has a consideabl afet on allowabe amsr lves. bcna fatWie srssn an critical. Thef=, it ia impodtant that blades be impgcwd carfully for maggnsl quai-
I. Mas force unbalance is controlled by horiz antally bak. cin the blades agaimst a masw blade within 0.002 in. times the bl•d• weight with a Eminmum toklence of 0.10 in.4b.
blades-qu obneor brand joints. In addition to vrqfmig thos attribts afcag amn* insei must imuoet that tho bla6de aus eownet dunancoally. Arfoail disiwon am LqW to anum that mak blad will prosm 4~t thrust perf•rman; an iwdivWhual performan= to*, maboth mpramtimie ad usaanamary. Aifl dwauioss ol individaa blaes also a rsad to emt amedyasmac balsam sinq a s of blades, Amta ualtya rifed byOi mfi aton of asub lot of fsqgpl, isauivig I l- ad chmaicd tot of tae. waM by ha•d mse on sackPon. P@-
cally bhadangi tim blades Wia a matw blade witmin 0.0 in. times the blade weight with a minimum tolaracm of 0-70 ia.4b (IWIL-P-26366). The vedical balance sould be accomplished fw two orhKooa bade anslar peaitis. To a 'hieve ma balnce, small wights we added, usuay a t bl•de root mimo Aerodyo"aw belanw of propeler balance usually a controlled by b aiwspvmWW airfoil she a by iestdlnthe blades with tir mr ne stationa within 02 dgS of esuh W . in mgay ifai os however. adeqas Control Cafnot be Schoerd
ty, su~rfec cndkitm tuA -- in Wtcef of s, ar-6hali
-
I
wem affwia
treths W& uh
eld aMWFORMOW rllig
2. Mom momatt unalo
wiho
sulecive gomu
l.
is contolled by ve~rti-
in thes cn,
the
5-73
I
•
-
weighted average blade angle error must be compared to a apacirwation maximum. The weighting factors for the an ro along the blade are based on the aerodynamic lOading fror a specific operari4 condition for either force or moment unbalance. Usually, correcting aerodynamic form nnbalama also will corre-t moment unbalanc satisfactorily. The weiShted average blade angle error may be ohtamied by manually averaging the weighted blade angle errcrs at various stations along the blade, or by using an automatice blade aerodynamic balancing machine, Hub and spinner unbalances also can influence the propeller balance. Even thouh the hub is atctically balbnced about its axi to 0.0(bS in. times its weight, and the spinner is both statically force balanced about itz axis to 0.0005 in. timus its weight and dynamicaly moment balanced, close dimensional control and indexin mut' be maintained to achieve good propeller balace. Critical factors iidlude oat-of-plane ard inplanc blade retention
Iovcall
*.->!
.
5&4.4.1 Helew Dw Steel suitable for one-piece hollow blades, or for the $par of Spar-h blades, arc low-alloy Aws eqwvalmnt to AISI 4350, vacuum mcltied, in both the 364o Rc hsrdnm and the 40-44 k hardness ranges. Thee steals must be protected from corrosive environments. The leading edge or the entire airfoil may be protected with eosion-resiatant coatings or olaiwn with lea-durable paint coatinP on the internal surfa•c. Impact dmage is a serious probibem for a one-piece hollow steel blade. Wall .thcknfa that may be adequate for carrying stiuctural loads may be thin enough to dent locally. The mrent reduction for local impact daner is the combined effect of h.e gouge stress concentration the local plastically deformed material, and the stress-raising action of the dent. Frequent impaction and local removal of gougf and tei plastically deforme surrounding material can be und to proct against the effaft of this type of damap. Anotlnmethod is to protect the
squarencas, axial positioning of the Wade, eccentriblade with a hard, damnrreistant plate Such as city and squareness of the hub retention on tse pronickel or chrome. Howo:r. the stsenth-reducin8 effC.e_ of the hap] ,ostine n the ,eA stru-tu....re tA peller shaft, and spinner mounting runout and tilt. Thc balarnc r,%.Aim-era of the hub and spinn be considered in the initial design. usually are met by removing material or addin,,,,\.JC balance weights. After the parts of a rrosraler are b•lanced RELATIVE "separatcly, the assblWd propeller is bahvi a WEIGHT statically in either the hou•iz.td or vertical position to 0.0003 in. time the proJle w bj- &"at &i balance weights to the hub. T"f-bladt should Ih at. .
SOLID ALUMINUM
fi,,Az balance to obutir, cruise flight angle darkia thisflight. the smoothest operetiao in
If additional balatwing is rquimd. it may be performed on the air'crraft dynarni.'ly by nameS systematic trial w¢,;tt methb or speial instrj-
'Er-
mcaatin,&L&-hat bepulse syichronizer unbialanceindicating (PSUI) unit or a vibration analyer. Because of differca in bUde augular position loading, and poi4ly in nacelk. system respo, m in ftlgt it may be mwcr.ary to supplement groud •.ymidamic bakadnc.1d 'S of ;k= prp• with infjlht dy-
b.la. nc
SE
SP
75%
STEEL SPAR AND FIBERGLAS SHELL'
.
minted substantial weight rvductius, a unmmaaiM in Fig. 3-47. Advaned *apoitie blad con•,udrticn
-- 74
H
•.••-•-n..
The continuing d&tvapat of new a -riala and coLstruction techiWqa, far propelle'i Wdes has perhas the potential for cvm mome isproMvin t. "Thepat qgrapks that fow deal with variou kirn. of Made maltiWs. Adtiomul information i,' containd in ANC.99.
AN
-d
ZZ-
T ITANIUM SPAR AND F IB.RGL&S SHELL F.w
$.47
.
assalde w-4 Wc*ittae~sa
-dM
-
A
S
Iean is %144in satVtCtoryV thniques for fabricating individual cowmponena from the taps. Brazing and diffusion bonding &,c two niroccase with gret potential that peoiaw the ctipbilaty of being dewcopc and monomies.1 manul'c-, In~to highl rqprouci turing -ows." Dramatic increases in stwoLth-toweight and sliffocso-vawtibt ratios over sither wat;4 or titanium arm possibie with the new advancedco pouites. When loa5Tht is wwaetially unidirectiontal with Amreinforcing fibers and the matrix material is o hiighly stressed, an epoxy matrix appears preferable because of its lighter %acigt.However, a nasal matrix offers hgher s®&h and stiffness, where noaded. MAlo. the allowable sts&eIL1 and duign roduki of .naterial with an epoxy mutrix noraniay miut W,ZdjuWte downward to ull,3w for moisture abaorpt;io ir. service and for time griadWa modulus dsware Wada continuous Cyclic: stresing.
When spa-gwe1 vnixastii is iiJ"d. with the spa mtn14r Vnd a sturrundas the majvw Ing shell of a 441cict rnatzrsiJ& tltt £2& is pvotacted from erosion and &iiapkcýdaub t',;b-rt heOlt Intermeal surfaces of the "s rm'u be pztscccss kbwever. (Mass cloth ot i n-splhwtv~d pkiw4ir rmatmrs5 can be used for the ac odynznk~ sge3l awlt bc:Wv to the spar with adhesiives. The iaatw and wxcrAial propcities of the shiell can be talored by selected orientatinu of Whe lay-up. Hollow titanium load-carrying spari are being developed, using SAI-4V alloy. Because of the excalhant corrosion resistance of ahaniw alloys in gSc*ral, no corrosion prot*Aiion woaquiitd. AM airfoil euvelops adhuuively boao4.) to the spar jrovidics both erosion and impact dsmaje pwte~tn. Thorophilugs that can dqrzds the lore no pfwLwr.W fa4ut uvragdh ottitaniwn mand be considered in this typt. of conatrwlion.
4
S4&45,C Fbs Maiculi vhMatuitt Riid wuztaha fowu materials with a daxuity Ln the rnaia~citicr (ild n mmf~ta bennodwiady;,,M of S-12 lbý/R cvrmnitly tit used as tk heanI'A wawd ,~43shav # W ct c k-uv nx kVCw b= an caidyd forv twiil Wý .uin. et& Aiwvb bilktA. Thv. rw"3.j n&a*1I1%4
54&441 Cg"a~o Won OUi
4
of the pout -jihrj:&v*L kjyx. Cz."A nadic LaE. ssd nSn kvnudiC.Twhaior~ f4x lay-tpdu~e4 in ttwW aloha, mi;"Az, wfNn.-n Ww thermal coaductivity ALAtii ~~an wtic to ad rercasns ~ ~~vtA~algUratios with adequate WsA4aw-v A~1.~ut~e rrr~ ~~oc~ai~ e4M'iev adlumiou, and satisactors-tc'sabebya poviiC sj~alskcnh ~rcnasha= Of Particular impcwianca in bta4 ~ ~ RtWtSizThShy. ww ~ ~ ~ o toC -w IVAq, ek-, "i eUSaxt&4 1S7 LU* winsale is Elip. ability of she. foams to beratim icatica fabeý"i ,..um of tM bac eke in unite o OMulet caxitw 01r DeeM -9 c*Atnt fitct S~q imU epasty MIPW.6r S4A4A Stvanlta, AdVtjh." and bwoa i4 cM. Jl6WVbitru 6UI) wUwrfl &;,&kki L~ tli . aa aOUs ~~~ ~~devloamwnu; wis .. Tegnatx w44 Fibeurn,, Vttlp ,,
-4
.
.
.
..
Av
.sasafils flat uhwo -~
Thc4ý b 4 % &q&&& t am"s t pWalak e 2AW ~Thuy c~o," bi- ma4uwsao
pksa of fibao"a ccaia~5( trocihl@itu . as ikuq*~ taLitusjs for me&AQ au.Aw TWc~ kv& &at *.auJ advvs P
itet6,
~N44WA "gVen
Vim Ia aftMO I wfl if tug SWIVAM?.. ftmiJQ& C%~rqý lwUia6Jtd(M4w 1ite4.. Mai'A @i~rA
@~ i"a o1wcecii-tw
t4OL~iaAahi hItfi.044 nd IL;tt
th~ba'tsA~y~wLIuOw
htloud&.a
2%~~rwv
WAOWSM4VL
gYUtJMe
1%
rLdi n4 &&IOz Ows be;iQ Wart,
jaiML.JMO am vw-wziz W"t i nile1* etba w tacVOf kiv - *-if Inc adt"Uvtn boqz4 p~~.vtt~ i~ ~~nhuc .i jAs ini im jý#itd edt sr proo is tkb km ca* ghd da4 cv ta4Jk isA t~cso www4imiuknm.a is Atkitnd, V#4a biJra k CtOVya4* CýAht~fr tMtEILA.Lk "144. FAflCA*E 1.1114-1 % tiC-* Mfl2.%&Otvi &QOPa %, Q>'Ar 4g~o btA~chzwp scon ts t zw-ýU aofpw &QM 4b~.a rajej lwt C4~ k~rw gbfUtntt C!. Mj1Q
&%b4pb S43
6us4atl~ k~k aid a%~o uwttaaw m&.iýs bwaAWc ia~~t~~tat1~U
k:.
*,
Iin-place
fsiatlh,- 4lou akJ
'I
f
bk~~~~adtn wiib mcýts k;.ss&L Akbkc
--
N'
k
C4
amp"s of piopelsr design, determination of blade fatigue strength must be basd upon both compr
E MEAN J'
fatigue tests of full-scal blades and vibra-
*hensive
,
-- T-rr_ -n--n1 WPLCIIIENS
12LT
tory blade stresses measured on the aircraft under conditions rcpresentative of service operations. The number of vibratory stress cycles accumulazed In the service life of a propeller is so great that
vib~ratory stiesmes below the endurance limits are
A
neoesary for most operating conditions. For instanc assuming a I P vibratory stm. and a rotational speed of 1S0 rpm, 9 x 10, cycles are accumulated in lOD hr of service. Some moimentary or intermittent operating conditions can occur in which stres amplitudes exceed the endurance limit. Each cycle of such stress uses up som" of the fatigue life of the propeller, and it is necessary to establish oons•rvatively that tht accumulaton of those cycles can be tolerated. Curreat practice is to apply Miner's Rule for cumulative fati•gc damage. Vibratory stress limits must be derived primarily from controlle laboratory tests for fullcwale propelie blades, supplemumted by sopcimen teats. Ground and flight measurement of propeller vibiatory sresseis during aircraft operafion is described in Chapter 8, AMCP 706-203. The instrumentation required, many of the co. ide•ations involved in planning and executing a vibratory stress survvy, apd the interpretation of results are included. Certain of
_ /
RILL SCALE
C...
"--
TO ... C,
-,
CYCLES to CACK
SCALE
DSEt-eCTON DLOG
SCALe
Fipre -
ladgue Strenh Diffemece Dutwee SpWms wad Ful-cale Teasm (mpreunted by the coefficient of dispersion; i.e., the ratio of the standard dcviation in fatigue strength to the mea'n fatigue strength, at a particular number of cycles) is gencrally in the ran'. of 5-10% for specimens, but may be as high as 15-20% for full-4cale component.
Trh6-
a.-
-t .l
.-
I..
..
nn..,.d
A-nA
upon p.evious tests on similar components, a.tiipated service loadings and environments, and related sevice experienwe. The various regions of the blade - tip, mid-blade. airfoil transtion, shank and retention -t midbt be aonider d as to thsir dhanks and
thes subjects are summarized in the paragraphs that
fabrication details, their steady and vibratory
follow, with emphasis on the interpretation of results. An example of the application of Miner's Rule is in-
loadings, and their emvironmental exposIres. The fqufreafnb for fatigue testing of propel blades arC similar to those for rotor blade dicusied in par. 5-4 and in Chapta 7, AMCP 706-203.
eluded by refernce. 01 @ (dw
E. dlmce U
Stmmm
T"U
54&i
Compiderable information about the fatigue u.rwuth of prooller blad materal an be obaiued from specimen tsw. A dtiscusion of types of testing, numb•r of samples, an statistical in•tnpratio of fatige tat dat can be fond in Rd. 57. StL&t2
hilse
Tels
p-m totig can womplamt, but never rqlw buing , of fu-sclet producti composesta. b of the S i fsni-" tioe dileamows and dieoa ia rss sate, fI-..ak comp-Lie.ts tend to have boh loer mesa 1[t'. snh•os sad a ItI
rtAe
scstter diam coetvvtioal laboratory sped-
mwn. A typical ilkmust
m oft
dWfwr m Is sbown in Fq 5-4, wim the mes ftigu etrenth of flwls coaponms at logWr umbom of cycle is about cm-half tha of the spamima. The data scatr
3-76
7-
-
r
----
T,--D
"M -I
H
no.
pedim" Ta 5-8.2.1 Airraft Teats The instrumentation requir4 for masuring blade vibratory streass in flight is described in Chaptr g, AMCP 7M6-203. In gcneral, survey tests are programmed to e-
compass all sigaisicat re"vice oprauing coitions, with adequate allowance for variability and, whm
possible, for rute cbha•e andlgrowth. The resut of a propeller vibratory strem survey customarily ar summarized in plots of vibratory streis agnin the most peminent variable - suc as propele qse, insped povwe or time. The curv usually are
selected to show the highe
stra
sampl of a gtr summary
-k
in th tp, the
mid-blade and th shank rWoa of the bla&. An
is shwn in Fig. 5-
49. A full sciof sudscurves will form the bei fatigu Iklieanaym.
for the
(
Of '
,
lS.C
at certain critical speeds in. r.,dom 8pound windlanvironment, or some instancs of stall flutter. In such
..
-
" \cases,
-
•newvers, "duranco
-o
I"shown
_._
it may be neonehary to avoid specific operating Sreions In order to assure structural Interity, For vonditions, such as takeff and inflight mawhere vibratory stres levels exceed the enlimit, the cumulative fatigue damae must be to result in an acceptable fatigue life for the in Apropeller. Fatigue life determination is discued Chapter 4, AMCP 706-201.
•5-9
ANTITORQUE ROTORS . Gin
S49.1 GENERAL The tail rotor of a single-rotor helicopter is designed to provide thrust for counteracting main
rotor torque at all flight conditions and to provide QM 4.•
Zo
.directional •that,
,1
So0 10
D
M INtNCAT-t
2W
3W C0
variable thrust for control in both the torque and antitorque directions. Yaw control is effected by variation and modulation of the collective pitch setting of the tail rotor. The collective pitch is controlled by control pedals and normally is adjusted so at the deign point hover condition, the pedals arc in a neutral position. For single-rotor helicopters with the advancing bladc on the right, (the conventioniui
AikýPUD V,
IUFULUi, Urhom'
s
.g:
~
u aee Mmlans To detrmine strucural integrity, the measured starus in various regions of the blade - after due Consideration of beckgroUnd data and allowance for
increases tail rotor (positive to the right) thrust, producing a left yaw. Likewise, the right pedal will producc a right yaw, with the tail rotor loing to lower values of thrust ai.d finally into "reverse thrust". The tail rotor design goal is to produce, with minimum power and weight, the thrust necesary to meet the control and antitorqu¢ requirements. Tail rotor requirements must be met without the occurrmce of any undesirable vibration, whirl, or shake
chavae
characterisics.
FW* 4.
-
TY~sa
Sin s
c r
must be compared with appropriate
material strength information. The strength data are -ued from full-scmae and ojieciamsemins. in choosin appropriate struegth data, the kvd of mean Wrw from davp computations is muflfieutly Accuin. For olpeating conditions consideed to be ementially commuon the srns leves mum be below the endurance limit. Normal cruise flight must be tr#Ated as a cotinuous opma•in condition, and, in ose inK-BaMWl atio rmal diab mid dsme dold be
wr~ed simark. A lhauther omelderaa
The tail rotor is designed for the most severe ambi&it conditions iic.udifl :l-sm-icop.c.,-"itia,-, tilude and tamperatumr, and the critical altitude for the engine. The maximum thrust that the tail rotor must provk• without blade stall is that required to counteract main rotor torque., while also providing th specified positive yaw acelation and overcoming tail rotor yroecopic preesmsion eects, in the maximum 4inc.fled crowind. Prmvmo mum be indudo to
counteract distuebmacso so&ase gusts. Coesidraties also m tbe givm to tbe tmlosm due to inatre-o useam om- mai6080M W u wth th 'Aertical ue Ba, main rowor, ad othr pMru Nuuue i is dmemi howhehr myo"Wsam i Kmm of tAhe ilnpar. wbm thdW ~ sunoma of dos pespulr *0K be Eapede bas shw~s tha if th, tai rotor th~a Poorly rMpg"ae ALa. WoMdiMe Whaere thel VWkseQUINUMMes At the W"ila los-upen oMiiM s aftm eablishing that
-h
toy stv -- tqgh ikbe low m owud - nmg become t lo vu eipesble vaiat s in cnrewuenm. Smi moifions mftk In~smoperedo.
(e4.,bowpisMs fyaw •, 100 aiWewisda minlma dim*, ealtitude) Ma ", d wrpirwe" fair foiwad ligh
easly will be ustheLed kk
vc.ss, ths -- 77
tail rotor thrust cap~b;' st;11 be an 'yzed under forward flight conditions, including critical 550 HP,,, -Rb (519 maneuvers. Without such verification. i! may be R),. *[lb (-) necessary to restrict the helicopter operational cnvo. lope or to redesign the tail rotor when deficiencies are discovered during subsequent Rlight test. + l':4' Prior to final definition of the maximum required TI, lb (5-20) tail rotor thrust, consideration should be given to in. lb (-) creased thrust requirements resulting from the inwhere creased crngine power available with engine growth. -masn moment of inertia of helicopter in Refs. 58 through 61 give additional data on tail sdug Spcf-yaw, Military no 8 ft, rotor design. There currently are SpcfX - distance ',rom centerline of main rotor to cations applicable .specificallly to the design of anticenterline of tail rotor, ft torque rotors. -'vaw acceleration, rad/sec' Q.,-, main rotor torque, ft-lb 5-9,2 TYPCIAL ANTITOEQUE ROTORS R-main rotor radius, It Tail rotors in current use employ from two to six (9R) ,- main rot or tip speed, fps * blades. However, there is no basis for limiting the number of blades. The blades arc retained iai a hub i t, ainrotor to.utrequread to providpe reqfire allow collective pitch change rangigis from positive to main rotoretorqon, an oprvdbrqie ngtvanlsTw-lddtail rotors L-aihy re Teti oo hut l The ailroto thustcalculeted in Eq. 5-20 is the net thrust. taking into account all interference losse bladeshv niiulhigs nete thre. r mre due to tho presence of vertical tail fins, flow field, and type, 53 commonly is uWe'to coto higes.apii dynamic effects. The rotor witi be designed to meet flappingth ontro thern magniude. n tal rotrs eurmn on nE.52 ttecii preconc required for banding moment reduction also tetrs can comenste e usdt fr te aeodyamicand cal boveiing temrtrature and altitude or th. critica centrifusal twisting moenste thates theroyai bandet engine. altitude. Normal forward Rlighrt condi arc centifual wisingmomnt hatdries he lad to of secondary importance since the vertical ions fin low collective pitch angles. The blades may be usually is designed to unload the tail rotor at tail cruite. retained in the hub by bearngjs, tension-torsion High-speed autorotations and rolling pullouts are straps, or clastomeric bearings. tociia xetos Thefr lads tilued otor ar muh stffe in In addition to producing the thrus: required. the all modes than those used for main rotors. For inti oo hl edsgndfres fcnrl onrl so aeo rotor shouldbe designed stance, in the torsionall mode, the combination of maniloroot.Th
"
a,,,
This. tierandr thantrhestiffness generalyis frvetors ih tabne
th hgerlaine masadhg (Ref.
59).mal
5-9.3 VAIL ROTOR DESIGN REQUIREMENTS Thetai roor hdlproucethe thrust noccsitry for helicopter yaw cnradthe attru our-
mnsof the main rotor. This thrust must be pro-
that a linear control will he obtained in all flight conditions iner thr tail rotor does not openatein the structural integrity, tail rotor design sAll conside scoasi pirobt~1s. A
onvitonmeet found with tail rotor hdlicoptcrs e.g., in theprsene f atal fninthe wakeoftemi rotor ats ideward velocities, near the ground, and whim the heiotrhsapositive or negative yaw Band pon he pnutmcmdond~os between the tail rotor and the rest of the helicopter (Fig. 5-50), the tail rowo tkrust T,, required is found frm. Eqs. 3-19 an-0wbX and R, umaured infeet&e5.Gaaur
a
K
5-9.4 INSTALLATION CONSIDERATIONS Durinj design of tho tail rotor. consideration shall be given to the actual installation on the vehicle in determbining the taquired thrust. The rotot may be designed to operate with the tail fin downstream (tractor configurcion), or with the fin upstream (pusher configuration). The rotational axis of the tail rotor also may be canted with repect to the fin to obtain a life component from the thrust vector. The locatiou of the rotor will have been selected to provide adequate clearance from the ground and other parts of the helicopter, and with provisions for the safety of ground personnel (see Chapter 13. AMCP 706-201).
"ALI4. Tractor C..iguratim In the tratctor configuration. the fin produces a blocksgjs that causes a thrust loss. Tests (Ref. 58) indicae. that the net thrust available to satisfy helicopter requirements may be estimated from T.
/I_07S
t
lb
7S A~
05-21)
/
T
-thrust
a.--
for control and antitorque, lb
_dk ~ce
by the
by a lag of the tip path plane with rhspect to the control Axis, which produces an equivalent to cyclic feathering. As a result, one side of the disk is loaded more highly than the other; if blade stall is encountered, the additionai precessonal momuent must be produced by the unstaled side. This effectively reduces the thrust capability of the tail rotor. The rotor blade :nust be sized to operatel at lift cotfli. cients below the value for sttall throughout the operating range. The increased loading caused by rotor precession must be provided for in this sizing. The effects of operation in a side wind alsc must be considered on the basis of a uniform variation of thr~ust with pedal position. When rotor-induced velocity approaches sideward velocity, the rotor will en. counter the vortex ring state. This characteristic, shown on Figl. 5-52 (fromn Ref. 58), giives undesirable flying qualities and generally is avoidtd by pilots. It is prcferable that the induced velocity (disk loading) be sufficiently high that the vortex ring state is not approached until sideward v~1ocity exceeds 35 kt. 5-9.4.4 Direction of'Rotatift When the helicopter is in rearward flight near the ground, the characteristics of a tail rotor installation Ufhtc10
total disk area, dimensionless
"5-.4.2 Pashe CoaliguratiO In the pusher installation, the production of thrust creates negative pressures on the fin and tail boom on the side adjacent to the rotor. The integral of these piressures over the affected are produces a for-ce that must be subtracted from the rotor thrust. This force can run as high as 20% of the tail rotor thrust, but can be redu~cedby increasing the ax-adistance between the rotor and the fin. At a distance corresponding to
VV M. W
tam
ýý
a
~Wa a
n.
aaa * .6WIS
WIi'hcihc
'wp" moving
-e
.
25. S/A
0.264 -201 .0 Q
4..VU
-15
used with caution.
UH1CC; S/A =0. 143 1o0~ ZUH-1C U7..
UH-1D
w
AH 1G G-
X_
effects. Therefore, the pusher configuration should be
o6"
produ~c
AHI1G
offect of distance for both tractor and pusher configurations 6a shown in Fig. 5-51, taken from Ref. 58. It is possble to design the pusher configuration with lower lo-sss than occur in the tractor installation. However, because of Dlow blockage, rotor performance is influenced to a greater degree by wind
S-.. peslia
forward
undesirable flying qualities. The ai rotor can encounter a large ground vortex produced by the main rotor, which causes nearly a 20% deciease in tail rotor
.
S/A =0. 136 ___b
Wheki the tail rotor is oparating at a yaw rat.., aI/ moment is required to precesa. the gyroscopic forme. This momeset is a functin of yaw ratc *, tail rotor_ angular velocity g,,andpolar moment of inertiaI1., and must bea produced by acrod 'namic forces applied 90 dog dhed of the direction ol precession ins the case of rotors with Rlapping blades. This is accomplished
__ 0.8
_ _
0.4 TRACTOR
Fl~um, 5-51.
%%
IUHI
0
0.4 PUSHER
0.8
Fin Seperatle. Distaaee/Rotor Radius 5.79
ramr
Ahm
an ad~m"'%W al
-mm o
3.amr~,i
t~-
u
nbotei
iG
f mb of__ -bw
AQ
-wea
4tdPwr
rwa vahsiwg tIN rolvdl trota. o f k trrn0 ise (Rd pow, #4 to the 3 an intnedyi cm imno himn rwkkr prseto ZXIWof the&;ltdM roqsirdc Hroa CJ~rtAiI IE "t3Un mch rolor rduction. v
ca~ mddhisd ta her4oyate, sti
ane I C E0d
i akeiulwsh coatovr1 ail roto
r
to~yat
dvs!;xxt1 :o vowst wich zbe bott4xx bkgk sovig
o
N
forwnd.¶ I.,~~
~
---
j14--
VNORVORTEX
LEFT
'64
AIGrI
2D 30 DIS& LOAD1NG T/A. psf
5-.SEml~E~tEq10 Possible diastie rdwsuiortk
tail rotor thratt 4im
40
SumM to tlcingiR3sS oa byt'ac tai r-jnor of h~ot exknatgae
(Jhith rwvihxat r~t~d air deoaity) should 1w atos~da .
nd by subvr'lqnot
tesing.
IMm"
I VG PAR t duo i u a alrtrwl pbykO:1O5 5-3TAIL Thecfa pt4~rhssi1 oto "nl 6140 ~ the. losc~nl, tip wgjmtL blade twist, sAnd airfoi vxstion. Etmaaa; tbkese csailncs inf)ucaa the iIaJwb wevight, 4i is =owary tW dtnnino "c~ inatoritatotpin ct&r to op~iiaiL the dcsign.
r
.*AS
j
5-a. Tal ar* Pnfnmmme,
Few Sied.
siderasrd (Egi? to the left. Tail rotor wsiciht sod boo'm size inamaae to disk loads decrease. Thus tjhj final selcton of tail rotor ViLk loadng wilt depend upon ai OvCJHrade-omd ff of required jpamr vemus tai rotor dis'uaer.
S-9.5. Tog RwnWT1# ed T1s~ rotos Aft d4M~imd 1t0 operat at tip speeds Of 54-tA1.2 TogCRw U"M L0144g 6W14= fpc. For a g.ven thrus requircmwt, taW ini t?4 5111* rtlir, the powerC required to rotors opv-ratiog &llowtip speeds will nad highe psmatt thrust depmads upon the dUk Ioadi4-, &urd t.Iiditti to obtain the requiroJ operath~g C,. The ibis on the 4iwtsr naoktld. As noe~d in Fig 5-53, leoý tip wo'j akw) hwuc~scs the torque, of dhe drive systra. i1,agc factor botl, increms the overe~l Wiijbhý the pCM,, F@CWEG0 £WCR*C with UJiCSmad &A Io*iur i.e., thrust bOG~iUJg lbjkg, *erw4acts- To of den antitvrqre utyatm. Highmer ti p speeds can result minirAt the powu; rcovirod. kv tdu kmdJOS am in blade itnakymic comnpmssblity losse wrth cot sn...we 1'.*Iwrn, as rmosd iip.-.3itPrwoovdiug ptwr loam, *bh oonsol forovis blad 4,36!mw 0hai ,, disk hwxL-jg be nsffackflily b11% to kIW .4. ercon prokams sand hiowe noine
I
A 88
YOLMt&WISJ Sins
than 35 kt dUiaiin
'
$-
Boris (Y-4. 5-54).
>
4
air
1W
doe
Ya in raw 4Eq.
90
I
Y*ASIR)
/SKma/n
'.'' sowdoad aped wd/an divId bar E". S&, aNudes 5422 ski hM ib blab Me. Pql Lw tiiota pom..msmsd to a qpdfedb muagouuimd is yaw. -"a
-rvd
labvsoerd to osbouud is 3btwht dwaureigs mud t. Wbepawfrwipee.lod diuiska (sad, 11stm paph PS sdumy) of fthtai9 rawn a d. Sals or kow-epeud Aqt sec Tb. twist neqire dffwiecy bseome pumu to nxmama PCysW. t du kmSWea HoWsova, wish WmM An .psk~ai rotor Wfd''Smai blade twist
70_
-u
-
ot. bu m -W -O hWOu7Wei sofwe a VAIL ROTOR TIP WPE.fps Vlgu 54C T1.S VftbS 10 ooRaw Nam
wow
_F
S4j. Nad. Husk er sof ITMe romw eoiiKy & eam Nc q3W* WeP C aMY WAY Nude chord &Wd had. numbber. Usaeuiy, s is boat t0 kip OhN vale Of asdvadurJ blade 00lidf ob-44l betweso 0M and 0.06 fo vrnon of Srui.a. lure said aselaioma. Tlw blad number b en he sany
I.between
)
,(AMC.
The in noldity rmqed, dopands qa tip spaS, disk loading. maxiuma thiust reqund atd the airfid ±toadun The chai. of fte aidaod affect th a-mmoperating CL. this directy intaad* the solidty. The product (ecb) of the reqwnC blade chord c4 Dm5tunes WOlad numt &i can be found froma Eq. 5-22. knowing the maximum opertig tip speed buiopter yaw rple, ran blockWdap ff, and thiugns:quin.ý
ncut;
6K c~ C.(B)'1wihonly CAC~p-
UA ui~+ ($22 38A ~ tvhcc 1 3BR~ b - number of tail rotor blades B - blade tip los fator, dimncnsionlcs c-e.*T&;'ive '!d ord, ft oissrtA pe1 tWaic, AUS-fl' - pola! rn~;of K m rutio ofn!ty*]teAdwqgthrug to " taii Lrito thrust. 6imensonlc A - tail totor radius, ft To -'tail rotor thrust to oompeoSawc fog main tokw torqv-., Thfor X - disteoce frewct ouerlina of msain rotor to geatedinc of Itol rotor It
/ ,"4 xkTQ "Are +
-+
4
4
4.
"ta
MieAh
whusu-
u80111
imlwsoveral Wction airfoi The coice of theW skeet a dotaMauium -m mW wfostaaacs nakdo hl the .awainvo apacWE lift .OdThimt Of 16 6"W
Ar 66 "AM&'
.Ivk
a
Wa~g p~cluig mow_-- charactensicL Many WJa rates we ACA 01012 and 0013 airoi atin whech h~av a pkchimg nmend of numtatialiy zero cv wr the usualhoratigw. naaeU tp =av and wgjht .f tail rotor are diecjy tpadmi~ upon The mauimum operatiftS lift Coadfkjd=; she nuv or cambere airfoil is bein cm.66=t The amo important chuaractaic of two camubued section cocukidee xuitable f£g tasil rotors ame skein MTable 5-S along with the ChbiAbCtUU&iIWi ofw gr veUC NACA cci:_ for cmsran The ?4ACA 23o1 airfoil is typkiJc! a clam of aim.toil sections wLeW camber is puninai
over !hM for-
war ation fthe airoil. Compared vith the syanmetricol akirfriil U'Ie aanUSa in an improved CmaWx9 a.~'s nrWin pitcbing mornnt coefficient Cm. JCs ACA 6M,412 is typice! of the taminnam-flow urlt(MACA 63.6", 65, sairi 66 strie) ralvsc to jJ -tWvi low values of mrinimum section dra coefficient C,,, by inuintt-mnoe of laminar flow over mutW. J thair MI:f8O;'; aiocs with pldgb-upcd lift and drvag charsrTeisticu flUe to th", fact that the cr-ýcrm it dis'Mbuts6 over thiw cv&ri chord, the pitchino&marinAv czmfficicnit isqcfesr, n-shown In Tabk 5-5. T.; vigrtw at C4,, am shown for cowmarigvs Ona10i ly. VWOW 'sa. . 4~aMig must include Wsrei opri';jng vsiu of RtiC Lid Reyno!ds n;A.bf-Aamrilua leadig 04t~ m~iougt . Fa7itu-x Iq qtry nom-oeq otrtcctiots tn sldail actioo 4 wihý .tul iu deficiet IU rotor capahhtý
r
;4W
-t;
a-21
2
by suitable deap mWd balacin of she roto "-.7 STRLJCTU1JAL CONSIDERATIONS Mlaul. aqatave (acsedowa) pitchin moumat can b. 54.7. S.Swul Druaiss Niluad. For campos by usin a avseisly stiff blade Withpt'00. nd y baancng te ~The plawnm:, of the ustual frqumncims .1 the taZ: with pndymmme sems (Frg 5-5.ing mAy blsewemle aft~ofrtri eattmne&J INICOCIS in insuring5 the MtiUCutdwIsaw srodsaw~ (ig-5-35 itmay emategrty of Ohe syifm. The dials of main ro,ýar 1 so th Ca.mposm of slwingS forclasa;s excitation must be considered. For ia. swodynamuic tbtaqpve arodmimi Onbedi dmAspu toontiue,two, %u'. a sixz rota ortu mayla she tail pitching momts. Howeve. when She CG asmve standit %e winsixg aio, corrpmW~- s ato ineto th aft Ike centrifugal ente~ring moment (tennus raicket 'Z4k At ra. Ae e.Atm-W mfeunyi efa) maybe inceasd and the 4mAuited nie. in sydlit also my be found. Asiaryofti smaui, moiest may 1W occu. Aother method OF sion sources is ashown in Table 5-6 (from Rde. 59). emo isby ug bom.In d pichm ham~g Guideline for ;Aacemnent. oft51wtal rotam natssal boo. haadl that pisthinbostmays by ingaor firequencies have be=n denelcped (Ref. 59) and are.reio- tookaio bWO my hOn ~ti SicopemI SooAtai datrin respnse vihaio modesn 15 Hz~. Sold becnidrd S-*A TAIL ROTOR PERFORMANCE 2. Thc na.&nal frmwicnce. of the tail rotor htotdG0 wku an bw hem rotor Once she sIlIW not be coincident with, nor Sn eo.w ptouiiity ?to, anyShn performusce can be eanimated for the zer velodexciting farce firriqtwiecie for steady-state optrating ty coadition by usin ate sme promdciuus as frw a with the excjta.. coifoa inc hovering main rowo (tee par. 3-2, AMCP 706-201). -
tion sources shows iri TaW.s 3-4-should be swa:idaS for at kni th form two modes.
S
* TABU:L-1 Additional daisof the placement &tthcs.' natural AERODYNAMIC CHA.RACWIRISTIS OF froquiencics m-ýiy be found in Rdf 59. SEVERAL AlIRFOIL S*.CTIONS, SUITABLE FkORruin ldn TA~kROTfl ADI2IYh inc"Auizon of' the firsa four lowest frequency modaý (au1-of-p!kanc and inplanc) is sufficient to etatady-stata rotor behavior. The aerody-
AA111Oil.
CI'4**reprtwait 1 58 0 1 78 -0015 1 67 -0071
_____
NACA 0012 NACA 2J)12 NAL(A 64 412
natnic blade loAds may be calculated by classical techniques,
I
MuCOMPE NOSE
with the local blade segment aerody-
narnic coefficient defined as a fraction of Mach numbe and angle of attack fer the c;Iuserating condition (see par. 3.2, AMCP 706-201). The efkcL. of induced velocity, Prra1ri rotat, velocity, fin intfr-
--
and elastic flct-i~sck velocity. as appropriae. must be included to obtain tist proper overall Ic.rence,
s.ATING U *AWENT
loads. TABLE 5-6 SUMMARY OF TAIL ROTOR _XIT
T~~~_
COMPONENJT
a
OF CF
[IO
____C
F8EOUENCIES
TSOURCE
NEAIEAIVSOTWK#Y.
IPITCHING
AEFEJDYNAMIC
/O&TSTEADY
UNBALANCE.
flb¶ilr
OUT-O)F-TRACK STATE 5XEDSYL-TEM
fi"
INPLANE lt
.4;EXCITATIN S-3. Com rtlemi for Negfuth Pucibchn mtammal With Cadvj Aq-e Aft ilal (C
IFIANSIENT FIXED SYSTEMINI
INPLANE
ibiW,
EX C;TATION
'Xefrt
BLADE MODE
I~
.,
OUI-OF-PLAN'E IP ~ N
OUT-OF-PLANE MIi H10 AE)ECTION ________
INPtAN* 1'x-o.PLN
4k
S41.73 Emb Sfiuua, Assi.el When blade sections. stiffnes, and man dlistributiavg have been selaqed and the externally distri-suted Iceds calculated, the neat task is to establish the
integrity of the blade. The internal strain
qmacy or tbs Loom. in torsi. was omiplid to the ch-ag in rotw thrum due to lhteral vulociy (6flet) was eleoauuaamd. The so that negative dqmpi c'hi'p ta tail rowo thrust due to the amotio is pro"60441ona tohashfila1ppiNg b &andlths pMb.*-fiw COUPhmq tan '1- Item WOfaed is this ma tat" aMW Live k was the best way to damp the sysuma. This type of problem can be avoided only through c"~efWl canskeato ofhe dhelicopser. cmiao Ucm
There ar. tkvo gencasiateigories of design loadin wonlitions c -c;sidercd in bin-kdesid-p. ultimate comditioms and !fligus coditions.'Ibe blade mus have an ultimate sirenth "4% greater tin the highest pea load santicpaWe durnng the lifetime -3( he sysiem. The bklr~e also rnuat have fatigue satength snjfflcient tc pre~vent a failurt 1= La alacmnatinj loads, I~~~Uatpence has shown that fatigue 'salnly is the more uaiticul design condition. Inma analysis a spectrum or atigue loads and tie:, expecte6nmrc focurn.*i the lifemtim oftkLet~c~' sdvlpd h sdrvifo !.tr;ir l nnucl tFL bulair'cý, climbs, T.%*s,;,' lostr. gadl ground-airgrotril c,,cla (sz.ý Chapter7 4. A MCP 706-20' and Chapi;' Ahife ;)-W203). The fvý, , stxcpnlh of the cuy.;paw,nýt ucencraily it
54-7.75 Fh~ase d DhwssMEu Tail rotors arm no scaled-down main rmuors Thur.. fore, fl'atter and divergence problems gasally an not as severe as for the main woanes. Usualy. the tai rotor Wlades arm much stiffe then those us main rotors due to fth opeasting eaviroomeat. Also, the relative inertia. as exprissed by Lack number, is of the order of 2.5 times that of' temiror.Fnly, aspect ratios of tai rotor WilJes are azuch lower. Although these fmactrs reduce the tendency toward flutter addivergence, problems with tail rotors have occurred, so the proper combination of 3 mad pitch linkASfiffnee Watut be iV,,' 11W InMeaddeam for the main rotor (par. 3-4) and the data in Ref. 59 mnay be used to determine the design details necessary to eliminate this problem
structural
(streWl distribution among the elements of the blade is wemputed: the resultant stres Levtls are compared with allowable iewels th~at test anid/or experience have shown will preclude failure during the life of the sys-
tr9A
cr
j
I ~~,
4,
.-
given at, i_1 '>tt 4 a S-NI ý-rivn dcneuive from test date. The rngnitudt ai~d shape of the t~rsrve art a function of material, stirs v ilichtaui'sW1, en Amonmental conditions, and magnitudt of conuurrent srejadl inrcss. The influence of steady stress on ftiti4,,;m strengt m.a y tý& considerrblc and m'jst be coansiderd in fatigiuw analysis. In tail rotor bK~ts, the eleady &za generally is equal to the altcrrt.narj sMtrcc, a t. reduces the allowable alternating strews by aprroxiniately 20%, depending on tix. mat~ia. As with main rotors, tail -otor blader dt'ign i! ""iterative process of developing a blade with adcqvcer aerodynamic, physical, and dynamic chaxactuýs!ics, and fatigue integrity. If, at any stape in the dra3ign process, an inadequacy is de'aced. the design process begins once again uintil all required chatacterustics are achieved. 5-9.7A4 Auroeastfizlty Problems due to acroelasticity have been en. couaterecl in the installation of tail rotors. Thesn problems have -esulted in undesirable flying qualities at bipb speeds. An example of susch a problem, termed "tail wagging", was encounterted on an expenimental helicopter. The rotor was mounted on a fin that projected atove the tail boom. A natural fre-
R'LFE!ENCES
I. A. Gc.'ow &.aiC. (G. Myers, Jr., ArrapynvWaic of,;he hrc'wcopter. MacMHilan Co.. NY, 1952. 22.HI. Ii. lijyto'1 and S. Kaizofi, Notate! Coniponseng of Inadua Velocity of a L~ftiq Rotor WihA A SoadO~ Disk Lioains. NACA TN 3690, Aptii 1956. 3. K. I- Licxltcu. "Srnrt Aspect of Convertible Aircrd:t Design". Jouweal of the Aercvsullcol Sclenires kApril 1956)Y 'I. I. X, W*ortsaun. und J. MI. Drees, flesign of Airfoils j') Rotors, presemted ax CAL/AV LADS Sympoi:9um on Aerodyranamics of Rotary-Wing and VIOL. Airc-raft, 1969. 5. 3. S ire .i, R. 0lson, *and %.J.Landgmebe, A At'avsctasnent jr Rotor irovrriqS Perfornnace Fredcein Methods, presznted at the 23rd Annual 1U,tional Meetiuaj of the American Helicopftz Sockiety, May l91,0. 6. W. Y' Tanner, Charts for Estimmring RotaryW1 -4 Perfonnance ifot Moerand au Hugh Fonrvtd Spruev, NASA Contractor Report CR-I 14, November 1964. 7 E. L. Brown and P. S. Schmidt, "The Effect of
Rotor-Pitzhing Velocity on Rotor-Lift Capa-
bility". Journal of the American Helicoptcar Society, 3, No. 4 (Octaobr 1963).
"531
tC. L. L~istpou aid M. ft. Mufhy, 'riy Qu~ms Conidrauosm 6n the Du~s aPA UN Ok of ibs kuayCtksw, Jant or~m AvAenca ltiacup U~y K$ No. Il(amary 199). t Nt Chii Ctoa,4 "Pitch-ka las~abihie of Hi~cap ur RtoWn", Jotwaal f the Auaican Hehcopwa Sockity. 3. K3. 3 (July 1958) 10. B.Sa. Bite, L6 L tflthnm, and It.0. Loamy, "-RantSuidin of t6e FIIuh.Lqs las""bhI of
flawy #S4f~fxclWd MaOil"uA AW Hdkispur Des W&6 kikgod Mos NACA I 31014, Wawbdiagii DC. Felbrsuy 1%7. 23. WI~laa K. Ocmewar.r ft~ysm&Iaup 4r ildiwpzar QeedW". Vbnaes. Prosl a pVO of th 13th A4n94al National MWAinG of the Aseeka HficepuW Sociaty, Wuhlwgotr, VC. May 3157. 24. Onepa W. Drkam, The Medeuiks IrasaWty ad EasY Aspesa %f AO&O" cw Nhladgd
stly Doetmi Disawtaicn, 1j2. Helicpter Socictv. C, No. 3 QJuy 3%.!). 11 Rt. L. UisptongM,-f, H. A 1 L Hoffmn, R.y 25. totert P. ColuNan, 7htwY qfS-ezcltf No.eC""e On nM~s f liked Rowr Mts. Aeroluskidty. A~dac~e-W*Aq, NaVkhkqt Co.. NACA AdvmnrMl Rawikled Aqonur 3029,1943. Reading. MA, 1955. Wt NACA Repot 1351. I. doiktotoclladsotiouaoW 1.M. 1. Yout,-"Ahsy a&d 1. E. Garrick, Nata,. qf the 26. T. Thecale Stability In Powcrd FW-iat", Journal of 1965). 13. K.)3. Hohcnannr and P. W. Heston, Jr., "Anochastic Instability of Tonwa-ally Rigi Holcopter Blad&'.V Journal of tie Amexian Hdioptcr No. 2 (April 196r,. Sc~4ety, . 12, -s. n -
A
It~1.
t5.
36.
0t.
lB.
19.
20.
21.
J.
MI. LOOtOU~,
.4
K.
W.
IRVIVC),
USC9
0ý
Om.
L.
Blankenship, -Computer Flight ToiKinS of Roaorcmar',". Journa of the American Holicatca Society, 30, No. 4 (Omhobr 19%5). Charlcm L. Livingston. Billy J. Sird. and lyce T. MeLarty, A Seabikiy ead Caxndo Prd-icun Mnhtidfjr Ile/4Oleoptr am Stappable Meftr A4kcraft. AADDL-TR-0 323 (in four Volumes). NK 0, Mytletad, A Tdvim, Mnhodof Ca'cu MOVin HeficoprrrMakd Deflecdo foraedk.4oinenai paper presented at the ASME Aviation Division Conference, Los Aragckes. CA, May 1947. r,. L. b3iankenship andi K. W~. Harvey, "A Digita Analysis for Helicopter Perfornnance and Rotor Blade Bending Monients", Journal of the American Helicortu Sodcity,?7, No. 4 (October 1962). G. L. Graham and Mi. 3. McGuigan, "A Sim plified Empirical Method faor Rotor Component Fatiguc Dcaigtu', Jou;nal oý the Amecrican Helicopter Society, 15, No. 2 (April 3970). Angel F. Madayag Ed., Metal Fatigue: Theory and Design. John Wily and Snns, NY, 1969. Carl E. Nord, FEudanin due Rr,4ieilfy of Fatisue Leaded Roiorcraýi Smnrcwvs. Proceciings of the 23rd AnmzLA National Meoting of the emlian14k4t? ociciy, Washington,ti, May 3967 M. L. Mil, et al, llelicoprocs - Cdcdsario. and
Deasgn. Valun /I. Vlbni" a" D.Vrwwc Sts. 2.bllhv, NASA-TT-F-S II, May 3968. 2.Robert P. Cowana sad Armild M. FuimAod, 5-N4
&%V"of eke Plum, Pnbla NACA TR S5, Was&hingtoe, DC, 1938. LMIWenay xplnA~M of APr P7. SMIUS PknaL Anm FbwE Mtekwsbsx Prtntdfiq of the LAS icsad National) Spcialkn -1 Me"tn on"VW Dynam I. 1 W 1-t... AW
MX'ma
%J.M%.um nnutaa
aaq
1908. 28. Jan MX D as. Acivelsflc Rotor fl..oa... and Niw-Stand, Aer modybaus. luternational Congrms of Sut~aci Aeronautics, NY, April W96. 29. Robert G. Loewy. -Reviewt of Rotary-Wing V/STOL Dynamic and Aeroelasti Problems", Journal of thm American Heicoptm Society, N4,I No. 3 (July 1969). 30. Norman D. Ham. and Maurice 1. Youns, "Torsione! Oscillation of Helicopter Blades Due to Stall", Journal of Aircraft, 3, No. 3 (May-ian 1966). 31. F. 0. Carts, "An Analysis of Stall Flutter Inntb~ily of Helicopter Rotor tiladas", Journal of the American Helicopte Socieiy, 12, No. 4 (Octoboex 3967). 32. Evan A. FraderJ~urh and Riclt~rd Mi. Segal, Model &-d FutI'Sceie Conyound Helicopter Resa&*A Procedings of tbc 21st Annmal National Meeting Of the Amcrican Hehicoptc Society, Washington, DC, May 1965. 33. 3. M. Drees and Mi. J. Mic~uigan, Hlgh-Spe~d Helicoprers o4K Ceapowscd in Maneuvrs and Gain, Yrocssdis of tbe 21rt Aninual National Meeting of tLM Aniuican Helicopter Society, Washingon., C May 1965. 34, K. W. Huwvy, U. L. baakmnsbip, and j. M. DrhIGs Aneoytlcu Laity of Hulicope, Goes ReaPOW at H40*Fo~wti Speak. Tit U-1, US Army Ablation Materia I-sbor ~soi,fort utall., VA. Suptmbtr 1fl.
K
35. K. 0. tAct and F. B. Gufafa.. (bAnf4 LaNot".o( LawjfAdwd Slattay De~jv iai4a ltiakeper RA~ in Fumed #~I&. NACA TH 2309, Wasahiagou DC, bLeah 195l. 36. A. Omaaow, Eymuams and Precalns fer
50. J. C. Houboft and G. Owoo.\ tW~man Eqssfl,awe fmuc. Chw~i:sma of Medsiaf 4 %wd Aw&V. a4d Train f Twined Naulfn basar Dliet. NACA TN 1346, Waebiqom. DC, Februar 195-F. Nannaiy CalwingftO Awre~pod (lar. S1. J. C. Ho.'h* Caspled 8adtq ad Twatae acunitaedcs #1 ft~r Rena. NACA TN 3747. Vtfewmuns Vj flwaV Rimahq DSwks w~for Waebhqiou DC, Otober 1956. Arbflwy Lna IAS Preprins Nc. 539, S7. Runul f.. luiquis, hddctzps Gun ktap w Janufty 1333. lWAdsdl Umsawdy Aen*%ade SUN EOfecu. 52. J. It. De~c and M. J. Turner, ropewPe YR 72-M, US Amy Air Mobility Retch anid lutuC, iou-san or the Actomaaical Scimi Developmnent Laboraloim. Eupia Dsrntuw;A. Ilk No. 6 (June 1949). Forn Eutsi, VA, May 1973. 53. Ii, A 1heertimd 4ppvacr to a Seed, cf .¶aE 38. John C. Kelngn and David L. Kidd, A Study 'if Flwer h.ob4ma Cotmell Ordajut Sboaoo of Tmvbdt-Poemvd aHmpelpr £4 flve Svaim lmnAAcronwtis: Eqinueriqg. AFO6it5 599 Juuu hdlly, Poomefixw of the I.t16N Anne &. Nat6ona 1961. hissing of the America sl~icopvcc So(ciety, 54. F.a. Cauta nd C. Niebach. Pat WA 4aMeIA, Washington, DC. April 1953. Imitabidoy at lila Fensuwr £puedr. Vdh. Ifl 39. SAE ARP-704, Heicopter Esgle/Reiwr Sysemx Smil Rlaaer. TR tý4-IC. US Anujt Aviation Ceeupuadibty. Society Of Automiotive Eaesrs, Ha4auiiiI Laoao~ Fort Evlik, VA, Iý'ruauy NY. Jane 191%2. 1969 40. Josph L. Peakowtki, Aaaaowagic Contirolof Frre is5. J. E. IM~er, The Sifta..' of Venom Poranwees. Turbine Engines, ASMdE Winter Annual Mccetin. lnduatq A~ich Nueesder,. en Puope &vDod 4 1. J. Shapilro, Pdflncles 9f Heilcopser, McGraw-Hill Book Company. luc., NY. I1W* 42. AFDMA Simaidards, Soclion No. 11. Aetehogs of
'IC#~IFN 3357, Waahi..gon, DC, 1955. 3* S. G. B~, 'Popui-er Ftialncing Probk~hi&s, SAE Jokircal, 53, No. I I (November E945). Ewiluawdq Load Rwiotjs futw RAtes Jkmrý 57. STPAý i A, A Gaakk for Fcuigue Testing aadi Ac AniFitinM igMaaofaaatrcrs AssociaS~wisekg! Analyxis of Fat/je Dta1c. Second Si tion In., Y.lion, American Scicty for Testing Matersul, 431. EetginteMVi M~ammaa SF-9O. Rolki Bering Corn1963. pony of Ameiica, West Trenton, NJ. 58. .t. Rt. Lyiin, ct a;.. "4Tail Rotor !)esi1 0, part I44. AFUMA Standards. Section No. '4, Maethods of Aroyaic" unlofhemrcnHlEwe/eec:kmiz" Rthic Bal 14una eans.An Loa Ameica Friction Loadrwig Maafacturc ~ A~inss cAnt:copter Society, 13, No. 4 (October 1971),. FricionBeaiuSMarufaturat sso~atoo, 59. P. W. R&.ke, et ag', 'Tail Rotor Designi, Part H2 NY. Dnmc,'ournal ofthe 45. P. Rt. Payne, Helicopnkr Dynamics an eiyAmerifcan ~~~- Struaursid Helicopter Society, IS, No. 4 (Octobecr nartsacs, Sir Isaac itMan and Sons, Ltd., LonwDnmco don, 1959, pp. 361-36-7. 46. SAE ARP-926, Aerospce Recommended Prac. 60. F. Robinson, "Cmrront rThnaS in Tail Rotowr lice De-sign Analysis PrceaAerrfo& adw Mode, Design', Journal of the American Helicopter Effects and Criticality Analysts jFMECA ), SocieS~~,3.N.4(ctb~ 7) ty of Automotive Enguneers NY. 6i. R. J. Houston andi C. Ei. K.- Mowr~s, "A Note on a Phenomenon Afl-cting Helicopter DirnctionaI 47. R. E. Pctersen, Ssresm Ccncernradion lesigr Fac. Control irt Rcirward Flin&t", Jo'anal of Amertors, John Wiley andi Sons, NY, ubao Hclkitipter Society 15, No. 4 fC-ýc~to-a 197e') 48. CAM 6, Rotorcraft A irwoothinecrs Nornoi Cote v2. W'. Weisusr and G. Kohler, ToM ioior Desige gory. Appendix A. Main Rotor Life I~ktermiwnnGuidc, Tk 13-99. US Army Air Mobility lion, Frderal Aviation Agency,41ashington. DC, kcscarch and Decvoprnicie Laborztorica, Revised January 13. 1%62. Januasy 1974. 4.J. P Den Ilartog, Mcs*ewcaf Vibratbm. Tirzd 63. N. N. B:fr 4 , "Rtzults of a Ta" Rotor fJ.0rcctionEditioi, Mrflraw-1i! lBoc-k Co., MV4, pp. 331. of4-RCUto,'tts:t". Journal of the American Hcli' owtc, Ite', K ... 2 (April192
5485
AAs" cAarr~a 6'-
FLIGHT CONTROL SUBSYSTEMt
7
64UTOFSVM3OLS U -. roIZ tsaapi
duritaivo of rolling momeo wade rqs to rai d"a ft4b/rad-c 4 dihedra ustability; duivatve OF realiai meWenft wiih rweplI to yaw angle. ft4lb/ia 1, - roll anwue saditivip, dmntiiutxw of moant MWa Vsqua to Conro .q~, ft-lb/ia.
aa aI li n'-'
mI"dt d
vh
w
erlniir sim,%"hstrm-n
t liw aimpty theiis S ,guz, sheikihas bawia stttuibbly , M W21 intomarunass cofaw Mated doptu SMCPt do -Ill into S k ceqon awulIo that mussthumam folylow quift ht[lo ie b lafluamo Of hudicogita Mobiity reqsdrwenss Visas
a
ggwt uAI h razwrs0 Sn
VA
~
mit.1 h
r.Ik/u
et.
No-yawast&Kbty; dsvaevccfyaswimgwnonmn N4 with rejajsct to yaw an&P, ft-b/rad N& -N YAW cn1trot asaaaasvitY; dMexnve Of Yawtug ntortueatwith arepec to control tDf4 ft~b/n a -number of pitch lPus. dc~ncnionless suj:nrnber of nonrmdundanz c'oapcocnta i 3 having the. failulio rate: F, ', a numrber af r5uvidsnt womponcmis having th.ý fuikerv rate P, a~ -number of norzrtundant cost piyAudnts
4-Ij. Poleof Dhpsbre Ty; kelly, the prelimnmary 4meign msafts insa drbanilion of the flight controls and a first e.W"Lrs of stabditlt augmcutauo:-i subsystem dwaaralristc teat arm bdcived to be sufficient to permit cortspbaacc of the licopter with the stebhisiy and earnzOl .ipcciFcto.Thcptjainminr dcakg dali. indu&4-con. trc' kinenaitics as firnacs on rotor blade or aaadynamic control sbarfiicc trawl, gciecrai anraregivnt of controlas, and mtdcbaaicsl &a-tx~s Tkc PAV' limia try dcsgnKvL% as a base Point fromn which
.a.atrc iolaumse jrat 2 aLnuiarc. at redundant comnponents
dcýAtvn akealswul arc proposcd for thc purjposc of amtpyovi'g ayste-in capsbilizy, relkitbilizy. m~aintainability, arnd cost. Thwu design alternatives then arm sub-
P W-q
o &
A -CR of oftaa No mcp itch dmpng duiuir ofsi rag-eft 6-1.1 DOEON MEtHO ae qse~d stabliy' duivativa of pisthieS hwcTn= aidrtrn .amrfacsur - &afte a IM ayes Hant with Muuze to crwanl vdocKaty; intqlrstoa of arirfmaze watrols, and uambiity amgft-lb/?Ps matatio ssbisyama - should ooudaac iterative ae jWiAt saabdity; doaiVa~im of paitkz% nka' and/or competitive trade-off studies. Thaci sadfis AV~ ampa viaý tpers 0o pitci ansit, $A-b/md will evaluate the purfoimace. cost. safey. roiabwilaty, - ladA coaziol auiatv-Ay: dcnvative of and maintrawnc charecteriatics of owe or sevarul pss"u* W~tib il L nIVec t onto contw @systems as they relae to the mission maquime- amagft-b/in. MM Mi sthtPRhabilty qmWatioa. refute i, -yam 'Sampiag dcrivativic of yawring H50to within this chapter. S
K
6-1 GENERAL
*
Z,
havirig
the failure raw~P2 fcilure rawe of uajlrqgatc of c~omponents, hr'nea~t. resrp vertica dtamping; derivative of verlica, force with rerpet to vertical velocity,
Z& vcst~ical control sensitivity; dcrivativc of %-erti"Ia forcte wih rcspect to contra;4 input, 0 - xjiafizWrwaton aout ý.yaxnýd
a -kocd, rotatoracl radise
totetd to A Uad esAiifgff trade-off Mtud) Cos maniaTypical considrratioais to' 1w. rivvi.wtd arc: 1.ywii The level of helicopter stability rcquirc-d. wvith or without augmcntation 2. The paramet'.r& that should be controlled, and 3. The automate4 ta~.s or tutopilot, (plot relief) functions that should !beprovided init
swell as toexrnldsubcs
S. Th(.use
of
snl,,da-
rmlihn.
- qZMP Imtrwfr40ac.redisMaccd systems; augmenetation avý'nuavr locaeo~i; in)l00 (and-I wetheraumu' w Corfroo ~mlcd/amc Wtorparalle to the pilot's6 iit puts
V
4eu
AMN 'LCsMW qau AgiaC WkWUe 90 eMWUWga frelsL A oflason cm a oaid a Vapi wtoo rom is p1- issAd MLOC-SI.
aquas10 6m83 SW
.-
rMW t oth LA1M
smaN. AND
a -
lU
-e Mk- de-w%
n
SS
04bi dmmigMa, hqrne"
A.1
MiL-HAI i imasdims Isw wbiA do sablty o do lbs huluer a" be wa mesd. lbe fcwqans addkitlud passders to be Ac~e hor em* awsud lms. qaphbmiis to.waasud MiSim t. Co baits mn lbs wi4l. keefL Mmd InWO ONOaAieS dka5 o aShir .MOWgh musefif in eMpurlee with aG&e pusMa. (CO linis. we.a (1110COi..OalPs week htr mosmmtsied bel-.
-%
2.Allowable rotor WeS variationa
3. Fmatral load wdlgpusio. " ")&Ys ias~ se dependw Won the- fa~ovs; aim, &he heasieshsomd Is ovahiasac throaghous
"aP MnadON of wdis do eaht y hbdnpew tees qma mn ph-m MqI aissali MW a wa.n~ be-.les mise Onimmbetm am pnaga fts Wediq. Uses MW P e dugrma twa f lki no"id ai -m bs qadmu as vont 14 e n.'I COrW"bee, ioke IkAa sAu8=4"" p'aw - evalma& of smu/mA*M.e inosfi prsar and uelabef of lbs Ayine qmAm pa@ fttw S~eOf 111s pcssyM. Kt 1118 daMS. 0@u6soMS ie tima for NpsWaaim of "i"Mus so amy gmenus problem liadicusd by simletias. A msioita. cupa4i of evaiseiia many avsa of lha cors ai etmisd i to te 1.
W-. Compliacs wit sabiky qecifmcstons A*oud be evvualneas and all the hates weighed a to
6-2
stability ckdarcteruans of the baskc, ~aaupmcnacd airframe should be essablishd. The helicopter sysvem i's liky to underg swany thanges in its anionic controls but only limted airframe d~.p duna *blf cce
OWatzd UdJin"e thý Pio'a A&iY to cO"N" dbe vaceadhimpta ofis hursc repnses to external daujten Of myal.Aw~ u rapidiy with thes amt the magninadec or tdo s95 which a steady attitude or trim is ujaiawm, ald t1K exis16nc Of any cecil~lationh or lightly damped to. Engineer work to quanstify the rtirg of quialities and dynamnic stability so that, they er t~e t ro vidruitur.: vehicA wpih afrnsmk stability and to makc predictions or comparions for existing helicoptesb. The latter includez prediction with sufficient accuracy to Supplement fliht ttsting activities and to expand the enajncee's uiosstanding of safety-$-flgbt topics.. , 6-2.1 CRITERIA AND METHO OF; ANALV'ha.S The following list presents the isicre signiricWii, topics relating to helicopter handling qualities: I. Contnal poweti, sensitivity, and intcruis couplin; 2. Inherent or augmenitod static stability and damping 3. Characteristic- ro,,ts 4. Type of automalat control sysW~a and vsntablas '*;oatol~d 5. Fatsz feel 6. htgnitude of respose 7. lWAswv of cmoaalc uampoaaat uou stAbilty.
TOLSactions. 4-12 AALnCA
6-LIANAYFICL TOLShandling
There ame two basic forms of matChmatcal reprcscts~on ssesin or veicl trm ad sabiitr bWialt1-pcfiurbatiofl equations and total-formcqut-' tions. Tie typical skazii-petnusbation equations noted in par. 6-2. A MOCP 'I4i-20 1, expedite the atecalinent of stability at one flight couditkiv, and can incorporate nonlinear control loops -tcadiay. These cauations also are adeptable to paremnctric studies wsing analog or digits! computerm. Total-force equations con.plctcly de~scribe tkt absolutt forces acing uponi the bdlicuptc. Howeve, tisty require & rather l&sIc codapiefielit or either anaog Or idita computCr 'equipm;Au for thei; aolu"sIo.i7s typte of solvtiof. is nwussary wheft invesigaiq larige variations in fli~gh. conditions ýic., speed, attitwfr ard lazgu "h sagA. Windj tutnne tesint, shou4 L..- conoufta. it] crier to reine the mathematal woe.ý. Insatnatioccoo_ comnin aodYasMic cisaurwactrm o, Se faqW
STADIJTY SFI1F ,
di
ATUON
ve F
MIL-4lM rfllw d ois kabs gadidiasý 6Kr w aabhahaqol buakk qlamaeliusal Cnyasaip. wis dos spirmmita, bow~r. dos W Imm~ -Sauil)P pee ~doci dom dymamc Nabiliy beaoh .m~u le in ebut -mh wW46 Am -s -tam na ,-
<1
eauupamuy to ouzpk 7 aixoss, oreams embe ran B9m wham.sh valmr ot axis musiaty i ptnluo SW tip W IWL wb. prrvsc - moosiis *w yp.. of hukewpue smwamsa *NOW device,. in auiar
hwhws -
U U'
Cosidam
pasnn ta. bq.MvW sdhubapm
msAim eSlv~ia~ma, a6.6 web meiheds of sadl
0*, - n
w
inmm ta pan$. apb kid"$n6.
6-iLl CSmb.
tawo ad tm~bg 6-4..I, Amid Ut0a*~. pi Omeats Jui Me~t suautao axntr Paf ams Sto'q -M nau is for z aa~~n..ra...... &#a .l aadvieaaalirts ahns(o.- Duncepiek kian and1pi WW to (Vk).'aist P1 -M
IPar. 1-
ar fISWbIaP
ftJ
wa
.wslb rclkauol Ruhamui
asa roan-COOLolw
dofOabserote prilmn yWny incang wish airpeed wand/or wbids cosfatantaa vmarition. A Seine iwtiraons assen with impeac to fth pick. yaw,. sod asutods cons.auk Ft. 64I i&an illawastiamt of typia for bs rCCl 6111s. Can" fuAOoA" ac~dbedh ho0 Cowin flyisg quialiy usafnrkticn exiss for naLSicapaa rouarraft. Tharvkere abs psocurug activity geneally specifes acme commpoW upof MILAI4SOI adMIL-F47fl £sfuwboa off nimupud or load disaribuaaioa iamoul; the lift
VidLa vwalsalon tachSt bowmAenas shown an abs 0fO9ýt 64.1.2 cbmacinilc b011111 chimns. Itiii oavemfl-t to Umtntf iMI faqweateod Any mode if response ds~ccabke and/or directy camsps. dut to saaaasiy aumnwueamatri s Mans ii tW plout theme saunab- Wdalb ws. catro~inblz by te plo shaft exhbit a level of 4yTh &-ay diiu arnp xnt mjnati for m~casksaility at kwnq--M to tatwupcaiudminMULNO; for V FL oiI lcumaiirma
k
1, 44 .4
tanaly.M4 S
as appitn
Thic 1as
tan. (u VFR. bu i . a wmaimuai requiremwnt for any
onkqiusitnni
sidn.
Vcnl.Al cornuiol senaSitivat) , bar- VeganWd iaolping Z, l~iCC-vit ure Fated in par. &ý5_3 AMCI 1%
£8)
-
-
20i Thew: pana-iocicn arc fkxer. by pc.znnanc.. zon-I sidtiaauow. unt,"c~ thc.ý as- samec. b) £411 Cm"atOn vertit.;t au$'tiilai.4z is pwouably controis. Ust due to Lhe preset. tienu to iicotlscrL hove. -nulud
*
~
-.
~
4
lOTAL, AVAILABLE -/ r
systems wlncn no.. Co01tr31lboops Weit rep .aa vet-II tki caloq YiCyAnd post. -in. In hovem, thc thruwste kv 0b3v*'4a comnla"As
"PAL.
xi Thur aba are rcquir~ianws for stabiizinga sud responac mroilra as are not undler the diroct control of the pl';ot, or ate controlled by himt only indirectly stxch so by charging flight conditions or system settiags. Lxamples irickide modes influencd by stabdlity augmerutation control loop coupling, potential scro-l rnechtnLa! resonance, and external sling load dynarnic: Thcsc modes should mieet the MIL.H48501 cr..cri&, for applicable VEIL or lFt conditions, except %-tw cw MIL4I 8501 allows divergence er does n 0 poirxy a criterion, some minimum tcnptablc; damping shiould be designated (e.g., 0.03 damping rtioh, which is the minimum for flutter in MIL-A-
vmisarmarnPiipaces a n..saxni upon the manaM autinatatonloop S& t, including ant~ pies ~*masnnt ;gornpenantAoth, bincuuc £ high 5M in-m dio.- ea'galu rat'., and, tkhcreiors asngic dia, plwansuat a a gIvas. tinew These em"in ant. constrw~s~c #Acontrol sensiuvity san dampng ;urc%to. the Dom, -all, anu yaw coratror xe in~p&. 6-3.1, AM _P 140-2011. :ilth-,qtg flM0 sawng by thoic criteria is not sufficient, sacnxq J *ektit ant~ysc.. this methaod does se cvi asa fir it estlante loi-ratm v~eaih*i -s trul luu scim.tha~rr4a % -4 for at-e4ud' alstroi loop, wL.4c..- im Kt e"Waln;d for they sp'...dcatior. co~pfiiicg ay asnag tansies
~
-
_
I
.-.....
--
-
is-r,-m11 lei
cquas to thc rc~iprocal jft& wleruasi dnn 'piugTisla lag, lot a cagL, helicopact istho6u,5!eW (4-. [ m:').TWi £ IoI.S cru,.gL ro. p1.AilO0P.iO1 IAiLfI woiklcof;c to be influenced hnl.ac.aiby~ a 50A a' gicatcr lacdktaonM afta. cosafls. For cmrtait lasva"ptt tyw IL0'u4.- uhi-rotot M hides). utnt roatA iL&os.tas. effective. Asrcu of Woin.j&n51 muflunt thiugwhiai tOe fl4hs, ý.&vajs. Thivn~oa-. u
AILERON NTRIBUTION CO~
COTIAION -
--
.-
-4
-
AIR$PtDV4 I
TycIlienoaterAa&FltirSiala tc'sAo
Mal.-ptaM*. atasma a ** psss & cs uni ti tSh rm to ckaev haMma makpituc, aqma. tofmoe-g .dsptabW~s. o"4 4,
W-SM
nVi
wqhGood cdwA s3e. Aposo
va 2f
b ea
StZ n toy doe hacuia ofqWau~ow. hie h.ft msloe m ak bofh aim dazjsuy~ (and rus of cj chasauusain uma la dpes.s of e ub%-csanoicre in of& the b' ecwi.is d ofs~d Cnwlio etw typeiu of "i chave m the nl fomauaus Jof rotkm h pn;lots le qmwac f decy modats (msww o cW~iliuy
ti aa"had
TheN
ides rou t or to eis yoi aeggw m
pmm ~
Wcaai
apmclno 3Ausaam of wrnqm hwn thoe dfmping arpme msp ks eto hom-icaltods os mfo iaise withasurcins maq*fr theboam dicimatn. Aa Dasicpw rvAaamts Ad sky ausae mk as ih" mudbosownd 4tb
ofv
napess, aThs. forehbul-
mpads botham idnalotoathoin forw B
hoV
roots ofderarely Tileasc~i
for
mu the ho
ehil oashtioe.
Rd-ight
nd
A caracteristic rook pact fo! an unaugnsented belicopev is shown in FSg 6-2. Yin venical scale i. the dape na sutSl 5cqu~Y or co pan of the root, dm left hal (uqatave samana) daesiptes te sabl reel parL of the root with the time to kolf amsipb.aw as scaled, and the rash. half (pou;idve
(Cocqwe Rating of 3-5) wiales the mbodal damviugis greAter than 0.3. and wilt kopt a vauem of 0.3 only whoa the frequency is thCae favs worabls for a pvea flight tea e4%C-utrhnt. There ia no tarsea value for the shout-persod natura f~raiucnq; t"a mvt be estabhlshd best an
Riantap mngnumao a umakawaa hi
8"uoauvnass RAM
'10066'"Mu
wits
uunyn.T5
aaanU %ii. a UWSc WC%"V
tiant to aouWablisp-uaatu a=mu.ud
limits, however, because mnaeanog the frequency
tfhU boamitay lane satiety Lot apanifkmaei4 L-ttcria. Systck damping rs-tac, sac i6thtr tWe undau claas r *aompv;. natural frcquauwcs of die plottd root.,. mo ubz-airiaalct. Tilt unaaiaped naka.. squecnq &~as t&kLmzaaitud. of the rawutu vattroa Exir onen.. aaLA or coaippacr soi.1aiuns oful't 4 equauons of moutin may bie I iouted c -. zely 4i. thai forma- so as to inGlioat systcnr. proxtm. Y Ca aupe: 4lncaoix limit. taperinieaLsi data Art eama~wd eafler uiaag a step res~ponse and extracting .be dampnz.g arnd frscjuenc- illy asumarag a second-ox;car systum,.
There eAi be noob*-6somtzc Ejc~ .~aracteriuti attributabk to poor speed sability (pttxgi3id mode). Any long aperiodic or krag-ptriud (gwrthan 10 sac) os.hatory divergecrim, Wallowa4K k, 111C Present MIL-H48S0l, compromise speed stability s.r.d thus p-aoc al. aidded workload ont the pilot., Sdic-t-fixcd qi *tistability demands that all roorti cxhi~rtstable de :oping. Helicopter lateraj-directional fitsscit y nit s-ww gencially exhbiti toi suheidenhne, aprvai, sahi
or by aaawloLj.nampatcr matcmung.
two or mart: frequtnci-s in helieu.ptcrs with lateraC
Tht seczzca..oroi
1by
2';TypicaiA
DL. it roll modes simls~tr
to
those of fikxed-wing aii-
CUa-. The spizai and Dutch roll modes may ccciii at
ploas are shown in F4. 6-2 1ci a range da nit xi ubility adgincntaticn.
aW5$*spd kitnlou.nver to cruiing. ip.i. Ir. ww'c tat alosinglor cavrn ul uns ampinuL axniue nap~ud.. * henctope se toumaie aia smle anvrge.. t~eintidi; rue.. As di!.. auaxpeed iiaz.cascs. At sing rcr,3t movem to the riahi and b-ew~mti. a divni s; wa r. wwnfl the; other oscillator> root becomes swaM! atas
R. dlsubssocaccx as ain ?qxtrio-i reapvnsc. with tilt bot co;nsia-i1 set primarily by tlic amount of roill m; As The hi,;hez the damping, tab showie~r the
dpaeun.i:~freqefi).
Mixt cons .ar. Tile reqJartsncnt for mininiman datmpn4 is sct by MIL-ti-50l, e.g for a typical, light ($Yl)Vehicle unicta 1FR ~oiiditions, the require-ica.ii~a C 3-se. tume conutawat. For cargo hell-
T~s.ac R-Aqaed Damnping Ma2t.2 The mA5pagb ji ltufan i for citner VFR or ILk
ir. Lý xh of I sec. Test results Iron. a nap-of-the-earth (11,41 t),.cai Of A wt-por. CArtconnaisseraice heli-
phatioiiaime.imataw=. Thi adnipngqrakV
Wee a"GWL ;A oliQVUA;
0ptz:r. the IUS 1peuficatiog allows time constants
-t
auu I
0 2a.
bts rollcniEnLfrmC3t
nd uppllletal ataortro- 4*stro rexmware
1. maial3,wel-do~wý eouar les ikey o *tk inRes. ad 3 Te vluý o rlldamping
F4
fl,
12
0.8
STABILITY BOUNDARY VFR-MIL-H-5501 SATISFACTORY THIS SIDE
t
4: >-0.
0.7
1.4
0.6
-1.6
0.5
2.0
~j-0.6
3I
RADIUS EQUALS UNDAMPED NATURAL FREQUENCY
z 0.0.4 -2.5
0.3 *
STBE Cos-'4 art
O
tA
HOVER
0.2
4.0 ,5.0
0.1
__10
pi06lkt
N 12D kt
-"3
-2
= v"co~n
".
+1•
REAL PART OF ROOT
I
STABLE STIM
d20
TOHAL AMLITDEsec TIMETO ALFAMPLTUD, Figur 6-2.
sc
10-.0 10.0 c
UN UN~ALE A
DOUBLE AMPLITUDE
Ota~racedestic Root Plot
can be asese more effectively by the roll subsididnce thmc constant. For missions requiring agility, the timid constant should be minimized, ccr,sistent with other factors such as pilot accelaertion environ-
loops needed in order to reduce this time constant. Because of"the obvious requirement to minimize time delays, any lags in roll control azctuation also should he reduced as much as is praticablec.
ment and lateral pilot induced oscillation (PIO). In the twt rfercnct-,ý the low roll time constant ws obtained with a ri$jd-rotor vehicle. However, the same response =n be obtained with a conventional, articulated rotor with stability augmentation. Root and trimient anialymt can indicate the augmenuttion
At speeds above 40 kt, other lateral modes of the helicopter are comparable to those of fixed-wing aircraft. Unfortunately, MIL-H-85O1 presents no VFR stability criteria. Poor, and even unstable, Dutch roll characteristics can occur with MIL-H-850 com-n pliance. P•or Dutch roll stability under conditions
m
-a
MM re r" Wsim N WANdi Oag polwuts fidr v hu-m .ean the pilot. The mpoiemessW
sid 1obe M":.aiaews M- Mam. (WIntaq the aft pyle.) i the principal aces of in-
and DS* rol ased in pur. 3.7.4A mad 3.?7.$ of Rdl. 4 shoeS be ueed inite. Thu an i6iwir io
pom
Saued-sing upedflcsiona, sod am. be ma with rodirsnairy nbilft arefsfao. Amy coWqn if Ow MWW gn44 sassmad .•"awednds (rdeoammwo•y a "ateral p1qcd'° rImNN to - a "qWiral-W mode) •omd as be i " by the pdlo.
for sms-rar voces.
6-2.1.3. Veim of husman Tb. ra of added Abit g eqm ussia uo Un ni 's moft Smd be eiewe thoougallyo m eroot pots.q"Mw jy w* sWiad to down.ism aaynuica pis, soin eoatas air-. so othe flbs conditis Ths iiwms~ata "1i1i3 hh0~e Abba, Snhhy• will help ave tmin aid coo by hadieia rieas of Th mvAagied (imbmra) ailky of th bhecreed suibly an roam wheo uablit too moocolpe aO be oaioodwWfly •o •e wih ziect wb SeM t"W 'Mn S Pwu&Stw WAS U Px plo formats imuoinas to dee amt awonlgendw saad fRgh uewvlape boits, evn shoog I. MSine hebuopsar a airuqass. ateritical groa is planned Thiis a smwey bWarv aaqecusio.m aitusion cam mae w which &aqansic as e wighlts and as bash etr•sm of thOCG roaoa failuen requjir ta
-"
to fly only with she minest
-wd
2. Pause hdoeul'er WsuM aispoeked. as i
lhaw 1.
vehicle abiliy. A oo-Ocliwaem and pwrfcsM ip trade-off exists banwe the a -ef e
but at the upper bom deisted for tLh sinna 3. Helicopte with stability aupagaioe for the
de•s of flighMt eo
conditions -
sys&
ophuii
on ad M fae
Itsms I &d 2.
dre of ishemt nauility provided ky tshe lsic hblcopes Oefignion. In efCting sad a tads-off
-•systemu
6-•2
Typ of CadW
.atdy. the lad of stability required under failure coo-
The types of nubiity sugumetation applicable to a
ditiomw ! be defiwd siance it may be ubjec to variaion with he flht control symem concepcua. For c ,, ie_ oa eL.a•a-
given design vary i the parameters they control and, Iurdoe in how the aid the pilot. Co o.... .
ships between augrmtation sophistiaion and inhen• stability: I. A reliable, multiredundant augmentation ca tolerate a significant degree of inheen~t in-
1. 2. 3. 4.
Rate Attitude and trim Altitude and/or airspeed hold Heading
stability. A vehicle so equipped, however, m,'y not be capable of flight with all augmentLion switched off. 2. A sigle-channel augm•,nit.on system will require inherent vehicle stah.'" ty suitable for cmwpliance with the VFR requi'mements of MIL-H-&SOl under augmentation system failure conditions. 3. A dual-channel system will require an inherent stability level somewhere between the first two cam. A hardover failure of one channel (while operating in the normal dual mode) can produce a vehicle response that causes the remaining operating channel to experience a saturation in opposition to the failure. Such a failure, with inherent airframe instability, results in a rapid divergent vehic!e response, whereas the vehicle response with increased positive inherent stability becomes slower and more easily controllable by the pilot, A method for improving the inherent longitudinai stability of single- or tilt-rotor vehicles is the addition of a horizontal tail; "differential delta three" (rotor blade flap-pitch coupling on the forward rotor)
5. Hover position 6. Special augmentation. Rate controls provide improved damping of all augmented dqrees of freedom (potentially six, including three linear and three angular). This type of augmentation aids the pilot in coping with the shortperiod resnonses. but does not prevent Ionhlterm roll drift and possibly a sluggish or unstable speed hold, even with a stable stick gradient. The addition of attitude loops and trim functions (such as lateral accelerometers and speed-hold loops) aids the pilot by providing long-term trim-speed hold and strong sideslip roll attitude and pitch control. The pilot then must provide only the power setting, altitude, and heading control. Finally, the use of altitude-hold and heading-hold removes the need for pilot input to the controls. Hover-position hold loops, and controls designed to impart stability to cxternal sling loads in forward flight, are examples of special augmentation controls.
can te used with tandem-rotor vehicles. Lateral-
6-2.1.4 Trashet ResponMe
directional stability improvement is afforded by the tail rotor and a vertical tail surface on single- or tilt-
The characteristic root analysis only partially describtc the handling qualities of a helicopter. The mag-
646
mMW
usldlimmbemcs I also mot be desierld.
For- p
casur
impsabs. cn
kdL Tale 6
seaskivity
m uirmess age Sim by MN5L-14501 for &Nbutthe
"eic aaL
Sineaomphmiua"
ýwt
w09
deani of stabil"t aqimsmmaai oao- which may ad to in abs eosaiit. This e(ba m bs overcame by the incprpasion of coNtr ol t hmd-Isad awamnsIL TlS aP uS ol 0 a-s um of ho-pi. mssaic fim Smn
-
mama so steral ditanmce, but also provide dsh de d astivisy to t in addition to apfltmmi. for oliwtudiaal aew r vol sensitvity. MEL-H-US I0pents a evamisso for manever sabiity. It stats that, following a qd sep inpu, the accusatwio and p-b rate a Awk be cwa downward and canvsrgmto no usm than 2ec alter the mtart oflte input. A bort time is deirad for attaining a abaEnd valut of nmeleratbiwospoci lly whse inanevering capabiity is estial to the muia . Anallses can indate methods of maximizingthis per-
-
aye1ser
ft.a
naý"ae
feed-forward and loop closu. '
%Nmwanminnc nfl
Such analysis should
include any significant structural dynamics. For xmmpl elasticity of the -oo mounting cause a major increase in the time required for the response to become concave downward (Ref. 7). "Italso is deirabit to analyze the tranient response to pulse imputs. This response will indicate the helicopter behavior in turbulent air. An asseasmet of the attitude time histories. snd the time interval required to rea•mire the initial trim flight condition with alternate stability augmentation schemes, is ing some form of pilot control reed-forward, the pulse should be inserted downstream cf the augmentation system so as to simulate a turbulence encounter ,c.nce pilot pulse imput into such a system is not equivalent to an atmospheric pulse. In addition to the transient response analyses noted, transients associated with the following maneuvers should be assessed with respect to SAS authority, control margins, and vehicle capability: 1. Jump takeoff 2. Rapid acceleration from hover to maximum level speed 3. Quick stop 4. Autorotation entry and recovery -5. Hovering turns 6. Pedal-fixed turn entries and recoveries ,7. Fixec ollective, constant spend turn entry and recovery, The characteristics of residual, limit-cycle as-
pres-s. a ast of SWcwae limis for
- by a dim
w.
Fig. 6-3 is a gmphical pImssanom of allowa
-*
iatfr naa 7%u WMuriMM 1 ninsit S*moish the bigh-lrequacwy houto ansuola ras sntablibs 26• ifrIsy rane and andisplaces sKablIb the likOW-frequec mi
-
611.9 O Fman e stability and contral requirements qcified our ivod for thUe hliaopa eqamenIt the design ob-
jecaives for the flighot arol ay~stis Therefore, each
potaet flight control systm dmgn muss be inveiated to iumm that ma doe not violau• tho operational rsqummirmsm. ODA&. a tability a& metMAtio
syste
will exhibit sadsactory per-
tonmnae for mmall-perturbaaion maneuvers or disturtanes, but will impai vskid. stability severely duein lag disturbances and atmosphei turbtmlem because of its rate-limiting or maturation aka...
k .. hnm e tt tabish the a operational limits of the flmit control
system through analysis and exp4riment. If the system limits vehicle stability, it must be reconfigured appropriately in order to assure; spedfrcation complianc. Flight conditions for which such studios must be made include high-rate vehick motions, regions where stability may be £cntstive or highly nonlinear, and turbulent air. The influence of flight control system sensor outputs upon vehick stability also must be considrrmd end monitored in order to assure proper design and a pitch angle sensor in a bi-4y axis coordinate system will in',oduce significant inter-axis coupling at large bank angles, because the gyro operates in an earthaxis coordinate system. The effect of atmospheric or self-induced turbulence upon helicopter stability and control requires significant attention. A vehicle that has satisfactory handling qualities in calm air may exhibit large attitude and rate response., including poor speed hold, in gusty air due to poor modal damping or to control loop nonlinearities. Response to atmospheric turbulence shall be evaluated over the entire flight onvelope. MIL-F-8785 provides critejia suitab', for gust analyses in cruising flight. Attention also shall be given to special operating requirements, such us an external cargo hookup, where gusts or wind shifts often increase the demand upon the position stability augmentation system. Many nonhslicopter rotorralft are subject to ata6-7
ANCP 7W~202
bility degradation from self-induced disturbances closn to the ground during hover or low-speed operations. This phenomenon, often termed "skittishncss". is believed to be caused by the rotor wake
failure but the vehicle characteristics must allow safe entry into this condition. During entry, the characteristics of the vehicle shall provide a reasonable pilot reaction time from the point of power failure to initial
reflecting from the ground and rcimpinging upon the fuselage or wings, or being rcingestod as rotor inflow. With tertain combinations of vehicle con-
corcective action (I soc minimum, 2 sec desired), and should permit a pilot of average ability to maintain control with adequate margin. Once stabilized in
figuration and flight operating variables, this re-
autorotation, the vehicle should be capable of mild
coto ytmis fietosprssktiheste desinermus hatsatifacorystablit insre
su-
mentation adcontrol margin levels arc produced. t TABL MAINIJM 6-. APLIUDESOF LIMIT-CYCLE OSCILLATIONS ________
with
________ ________vchiý:les
OSCILLATIONi 414k
NITTC -IF-8786
L,.IIEAR ACCF'.ERAT!ONS
ThIý,J"1
I
85 1,.1-949
M114-876
MIT T-9
AEA
00 VERTCAL-n
proverrents.
5 and 6 describe the difficulties of obtaining a
/
SAA2
/
~
auturotation.
Mit 4 -C7R
................
.-
Interaction of stability augmentation system Magnitude of control trimi change resulting coilective pitch rcoucioao riC4Cb~d~y to C1isiw
ting the average pilot's ability to efltct the autorotation maneuver, and of defining potential im-
'DiSIGAD HLIý_OF IR F%'I"ENIRefs. FO PUE
I
Bufieting due to wing wake
.00
12 9
OSIAIN
II-I
3. Flapping and blade clearance with reapcect to the
4. 5. 6. irom
USALI-54
M!L
fixecd wings
2. Restriction due to blade stress limits
AT[ýNS ANGui AR fAT
-OSCIL[
AXS
falblwaaevlu.including that needed to maintan hdralicand electrical power, during entry into autorotation. Other factors to be reviewed in connection with thiis maneuver are: 1. Margin of control power available to overcome disturbances, especially near zero load factor and for
~\ ~
Figure 6-3. Allowable Pitch Control System Residual Oscillations 6-2.2 AIJTOROTATION ENTRY Most low-disk-loading helicopters have the capa. bility for stabilized autorotation in the event of power
satisfactory time delay in the event of total power failure at an airspeed near 200 kt. With dual engine installations - where the probability of sudden, simultaneous (less than 2 scc) failures is extremely remote - it may be reasonable to consider only singlec-ngirie failures.
6-2.3 SYSTEM FAILUJRES When a stabilitv augmentation system (SAS) is used - whiether it is electronic, fluidic, or mechanical the potential hazard of a hard-over component ~failure exists. MIL-H--8501 requires that the pilot be able to delay a corrective control input for 3 sec without the response exceeding an ar~gtiar rate of ~0 deg/sec or a ±0.5 g change in normal acceleration. The influence of flight conditions and CG position upon the severity of the response to an SAS failure should bc reviewed. Frequently, an aft CG position :oupled with flight operation near the blade or rotor limits is the most critical situation. Stress levels upon recovery ftrom a failure may be an additional factor in the ability to satisfy the failure requirements. Responses to SAS failures can be reduced in magnitude by the following meahods.-
AMCP 706-202 I. Reduction in augmentation system authority 2. Increase in inherent airframe stability 3. Multichannel redundant systems. Selection from among these methods during the design process is done after due consideration of the other flight control system requirements such as per formance (especially the authority needed to meet gust and maneuvering requirements), reliability, maintainability, and cost. Failure-effect studies, which note the consequence of each component failure, should be conducted in an organized manner. These must identify: I. Any failure that cannot be tolerated, such as an oscillation due to loss of feedback 2. Any compromise in control margin 3. Failure causing multiaxis response too difficult for the pilot to control 4. Ability of the pilot to switch out failures 5. Consequence of subsequent failures. As an example of Item 5, after the first failure in a dual system, the remaining system must niect the failure criteria or the flight envelope must be restricted so as to meet the failure requirements. Pilot-in-the-loop simulation is a valuable tool for
pushrods, to both the swashplate and the blade pitch arms. Cyclic pitch input to the blades is the sum of pilot control input and stabilizer bar teetering motion. Viscous dampers, connected from the stabilizer bar to the rotor shaft, control the rate at which the plane of rotation of the bar and rotor follows or lags the tilt of the roter shaft. This lag in tilting stabilizes or damps the helicopter pitch and roll motion. Additional data may be obtained from Ref. 7, and pars. 6-2.4.3.2 and 6-4.2.1 of AMCP 706-201. 6-3.1.2 Hiller Servo Rotor Another early mechanical SAS is the Hiller servo rotor. The two-bladed, universally mounted, underslung rotor has a gyro bar fastened to the hub at right angles to the blades. On each end of the gyro bar is a short paddle blade with airfoil cross section, whose pitch is controlled cyclically by the swashplate. Cyclic pitch imparted to the servo rotor tilts its plane of rotation, resulting in a cyclic pitch input to the main rotor blades. Stabilization results from the lag in the response of' the servo rotor to tilting of the rotor shaft, and the consequent pitch and roll damping due to the lagged response of the main rotor. Additional ur u
ubtainud in zci. a and pars. 6-2.4.3.
.......
U4*
impact upon detail system design.
and 6-4.2.1 of AMCP 706-201.
6-3
STABILITY AUGMENTATION SYSTEMS GESAS's.
Par. 6-2 contains numerous iz-fercnccs to stability augmentation systems (SAS). Owing to the inherently poor stability of a helicopter rotor, satisfactory flying qualities have been achieved in many cases by altering the inherent characteristics artificially Such techniques are celled mechanical stability augmentation. The pilot workload associated with early helicopters was very heavy. The handling qualities requitiments of MIL-I1-8501 have been developed not only to reduce this workload, but also to increase the mission capability of the helicopter. The result, however, is that it is virtually impossible to satisfy thrse requirements without modifying the in. herent characteristics of the helicopter with a rather sophisticated SAS. 6-3.1.1 Bell Stablllzr Bar Perhaps the earliest mechanical SAS is the Bell stabilizer bar. A bar with weights on the ends is mounted pivotally upon the rotor shaft at right angles to the two-bladed, teetering rotor (Fig. 5-7). Mixing levers are connected to the bar, and through
111n1y
6-3.1.3 Mechanical Gyro Refs. 9, 10, and II discuss two applications of intermediate-size (10-15 Ib) gyros in mechianic&l The first, produced by Cessna, is a rete gyro that is connected mechanically in series with the input from the pilot control stick to the control boost actuator. This system acts primarily to damp roll motions. The second, the "Dynagyro" by Dynasciences, is a two-axis, hydraulically driven unit with rotating damping arranged sn as to align the gyro wheel slowly with its mounting reference in the fuselage. Outputs of the gyro, i.e., its pitch and roll displacements relative to its mounting, are fed into hydraulic boost actuators that are connected in series with the pilot's cyclic pitch boost actuators in the respective directions. This design is similar in principle to the Bell stabilizing bar, except that blade pitching moments are prevented from feeding back into the gyro. The Dynasciences SAS also hicludes a hydraulically driven rate gyro mechanically coupild into the hydraulic boost actuator that controls tail rotor collective pitch. This provides yaw damping. No electrical power is required in either system. Ref. 12 describes an all-mtchanical yaw rate gyro for single-rotor hel.copters. The gyro, located at and driven by the tail rotor, tilts about a longitudinal axis in response to a yawing rate of the helicopter, and 6-9
AMCP 706-202 phase to reduce the d,:flection (Ref. 17). This concept has not yet been developed fully.
mechanically changes tail rotor collective. Rudder pedal displacement moves the reference poiat of the gyro centering spring, biasing the system for turns.
6-3.2 CRITERIA FOR SELECTION *S-3.2.1 Augmentation Requirements it is virtually impossible for a helicopter to comply with the handling quality requirements of MIL-H8501 without some type of SAS. Selection of the type of system to be installed requires evaluation of the deficiencies of the unaugmented, or inhereny, characteristics. The evaluation criteria include both the specification requirements and the requirements imposed by the missions assigned to the helicopter. Refs. 18 and 19 discuss the tailoring of helicopter handling qualities to mission requirements exceeding those set forth ip MIL-H-8501. Par. 6-3.1, AMCP 706-201, presents recommendations for contrcl power and damping. High-performance attack and troop support helicopters require high control power in order to achieve the necessary maneuverability. Good damp. ing in rol!, pitch, and yzw also is required in order to prevent the helicopter from being oversensitive and diicult to hoid in a given attitude. Furthermore, helicopters become more divergent at very high speeds, with the result that speed compensation of the stabilization system may be required. By means of simulation studies with alternate helicupter/SAS combinations, it is possible to determine a range of gains for the SAS that will cover the extrcmes of opetational requirements. During flight test of SAS prototypes, adjustment capability can be provided by means of calibrated potentiometers or resistors (decade boxes). Final values for system gains should be baqed upon adjustments made under actual flight conditions duplicating those of the required mission. The test program also will establish whether or not the gains can be constant, or if they must vary with flight speed, gross weight, or any other parameter. For a smail obsevation helicopter, the requirements of MIL-H-8501 generally are adequate, and the simplest mechanical SAS may be sufficient to meet them.
6-3.1.4 Lockheed Coutrol Gyro In later versions of the Lockheed control gyro, a gyro bar, consisting of as many arms as the rotor has blades, is mounted universally upon the rotor shaft above the rotor hub. Pitch links connect each arm to a pitch anm cn the following blade. Push rods also connect each arm to a point directly below on the swashplate. Sp,-ing capsules in the linkage between the swashplate and the control stick enable the pilot to exert a moment upon the swashplate. This moment is proportional to stick displacement, and is transferred to the control gyro, which precesses in the appropriate direction 90 deg of rotor rotation later. The tilt of the gyro results in an input of cyclic pitch to the main rotor blades. Rotor tilt and fuselage tilt follow because of the relatively high flapping natural frequency of the hingeless blades. Additional data about this system may be obtained in Ref. 13 and in per. 62.4.3.2, AMCP 706-201. 6-3.1.5
Electrohydraulic SAS
In order to achieve acceptable handling qualities, many helicopters use electrically driven and sensed rate gyros to measure rates of pitch, roll, and yaw (Ref. 14). These rate signals are amplified, shaped, cross-coupled where appropriate, and fed into lcctrohydraulic servo actuators in series with the conventional control boost actuators. 6-3.1.6 Fluidic and Hydrofluldic SAS The fluidic SAS, which is operated by air or liquid, is analogous to tke electrohydraulic SAS and may be substituted for it (Rcfs. 15 and 16). The fluidic SAS, with specially developed angular rate sensors having no moving parts and with integrated circuits having n3 external plumbing, offers advances in reliability and significant savings in cost and weight. However, it represents an advanced state-of-the-art, and it still may suffer from problems such a3 leakage, temperature sensitivity, and nuii shift of the sensors.
6-3.2.2 Helicopter Size The gcneral category of SAS, mechanical or power-assisted, to be used is dctermined by helicopter size, Only the smallest helicopters can use allmechanical systems, because the rotor feedback forces that the SAS must overcome are corrcspondingly small, In helicopters with power-operated controls (par. 6-4), the SAS need not operate directly upon the rotor but can operate at a much lower force levcl in the control system below the power actuatorr
6-3.1.7 Fsp'lup g Moment Feedback Rigid-rotor helicopters exhibit strong noseup pitching moments with an increase in speed or in upward gust encounters. One method of counteracting this tendency is to sense the pylon bending moment rind to apply cyclic pitch in such a direction as to reduce the moment. If the pylon is flexible, its deflection due to rotor moment can be connected mech-mically into the cyclic pitch loop at the proper 6-10 I
".L•
- W '- ,
,•,-
-,,.
. ..
.
.•
. .
.:
..• . . ._
: ,•
:.
. .
"
• ••
• .
.••
:=
•
. :
AMCP 705-202 (between the pilot's stick and the actuaton). In practice, if both hydraulic .ad electrical power are available, the SAS gytos arm made as small as possible and their output signals are amplified (electrically and/or hydraulically) to the power level required to provide inputs to the control actuators. Dual or triple electrcal SAS'i can be provided below the final rotor control actuator with less weight than a single mechanical system.
)
-
6-3.2.3 Type of Rtor System It is possibk to use the Bell stabilL-er bar or the Hiller servo rotor with rotor systems having more than two blades (Ref. 20). However, some of the obxcure refinements or kinematic relationships necessary for the success of the system may be overlooked. For example, the orientation of the gimbal pivots on the Bell rotor is critical in order to prevent driving torque from acting about the feathering axis. In the Hiller system, the amplitude and phase of the feedback of blade flapping into the cyclic pitch control of the servo rotor paddles are very important to the effectiveness of the system, The Lockheed control gyro, which is processed by forces applied thiough springa, is applicable only to hingelems rotors or those with an equally high flapjoing natural frequency. In order for a flapping or teetering rotor to exert a momeot upon the fuselage, it would have tc, tilt relative to the rIotr shaft. To sustain this tilt, the control gyro would havc to be tilted by an equal or greater amount, depending upon the b linkage ratio. This control gyro tilt, 90 deg out of phase with the swashplate tilt, would alter the phase of the maximum spring-applied force upon the control gyro, causing it to nutate toward the saahplate tilt, and eventually to line up parallel to the swashplate. In the case of hingeless rotors, with their high control power, the required amplitude and/or duration of control gyro tilt are too small to permit any noticenable gyro prew ione Inigeneral, whenever a rotor-mounted SAS is modified from its origin form, an extensive pro.... grim of developmental and qualification testing is necessary. The internally mounted electronic, hydraulic, or fluidic SAS's, which are more flexible and less dependent upon rotor dynamics, are more adaptable to any rotor system. The electronic SAS gains are adjustable individually in pitch, ro!l, and yaw directions; can W. made variable with airspeed; and can be cro.•-coupled if eesired to compensate for adverse airframe crou-coupling, H emode. 6-3.2.4 Helicopter Coafgraion -, ingle'rotor helicopters can be equipped with any
type of SAS that is compatible with helicopter size ani the type of rotor used. This is because pitch and roll attitudes both arc controlled by cyclic pitch inputs to the main rotor. The yaw SAS, if used, operates by controlling the collective pitch of the tail rotor. Thus, each SAS input to the helicopter control system i3 independent. On the other hand, tandem-rotor helicopters obtrin longitudinal control by use of differential collective pitch of the two rotors. Obviously, any longitudinal SAS will be required to change the thrust of one or both rotors. Rotor-located, mechanical SAS's - such as those that are used by Bell, Hiller, and Lockheed and that affect only cyclic pitch - are not adaptable readily to tandem-rotor helicopters. The necessity for mixinL all controls from the cockpit of a tandem helicopter before they are impressed upon the rotor makes it more straight-forward to introduce SAS control inputs in series with the cockpit controls before mixing. However, in large helicopters that contain many linkagas in the control system, even normal amounts of play in these linkages may detract from SAS performance. Therefore, the SAS output signals for the respective axes should be mixed eicatricaiiy in the same mannvi and ippici tiOrt a" arc the mechanical controls. Then the SAS control inputs may be introdu~xd at the input to the upper rotor control actuators. 6-3,2.b Suppression of Structurl and Rotor Mode Responses, Vibrations, or Gusts Helicopters whose blades have an inplane natural frequency below the rotor speed consequently have high response to horizontal pylon forms at frequenhies of rotor speed plus lag frequency ( wt). - If que.c.t aorotor speedpus agvifrateon either flexible or rigid body, has a natural frequency that is near the aforemnntioned sum or difference, there will be a tendency for annoying, large, transient reGponses to gusts or sudden lateral control motions. In the case of resonan=- at the difference frequency (01 = Wr), selfcited destructive oscillations can occur in the Wir or on the round. In some cases, the SAS roll axis ha& coupled with the r-sonance and aggravated it. Thus, steps rhall be taken either to eliminate SAS res[ onse to the mode or to make use of the SA3 in st ipressing it Unless a epecial design effort is mad,, the total lag of SA, sensors, signal shaping, and actuators at the high frequencies of the transient oscillations is liable to shift phase respons, into a region that caurts divergence rather than attenuation of the It may be necessary to install separate sensors, filtered to respond only to the pertinent fhequuncy and then phase-adjusted so that the final SAS output 6-11
-ý
A'
7ý,'
_____-I....
AMCP 706-202 is at the proper phase. Both the SAS and the entire control system must retpond to this frequency. Ref. 21 discusses the theory that n-per-rev vibrations may be reduced considerably by suitably phased control inputs of the same frequency. This type of vibration suppression requires large amounts of power, and shortens the life of the control system considerably. If such suppression is to be used, the control system and SAS frequency responses must be approximately 15-20 Hz. When dual SAS actuators are inserted in vertical linkage, the mass of the actuators may induce small control motions in response to vertical accelerations. In a specific case involving the collutive pitch lever with friction lock disengaged, the weight of the pilot's arm coupled with the vertical motion of the hellcopter produced a sustained oscillation. Mass balancing of the control linkage and/or the use cf viscous dampers are methods of curing these oscillations. Gust alleviation by means of control inputs responsive to gust-sensing instruments still is undeveloped. Closely allied to gust alleviation is airframe load limitation by control velocity restriction. However, the cos•trol requirements of the two tend to conrlict iocau-. gust aaieviution requirv-x rapid conStrol rcspons.ý. Becaue both programs have as their objective the reduction of airframe loads, the gust alleviation system should perfonm the duties of both. Ultimateiy, the SAS will include the fun~ctions of gust alleviation and load limiting in addition to flyingquality improvement. 33SSRELIABILITY The expr-ssion for the fa.lure rate P, of the aggregate of components in a system, as shown in Ref. 22, is PA = ni PI + / -'\l 1+ n2P2 +(!L2
!"2, hr-' (6-1)
where nr I
rtLumber of nomcdundant ct ;nponents having the failure rate P, -number redundant components having the failureof rate P,
- number of nonredundant components having the failure rate P2 n'2 - number of redundant components having the failure rate P2 A mechanical SAS having fewer than a dozen parts, all of which have a very low failure rate, is u!trUreliable compared with an electro-hydraulic SAS with hundreds of parts. Oa the other hanm, the weight penalty of providing redundancy in critical parts of the electrohydraulic system ii not great. As seen in 6-12 n2
r-~-
Eq. 6-1, the sggregatc failure rate is highly dopendent largely upon the at-iount of redundancy, given that the design insures that failure of one medundant component does not affeW the operation of the other. 6-3.3.1 Safety The designer must be cognizant of the influences of the inherent stability level of the helicoptcr and its SAS performance upon flight safety. Added stability margins can improve safety during night flying, or during limited-visibility situations caused by the presence of dust or snow ,louds. During such operation-;, the provision of improved stability levels allows the oilot to c¢ncentrate lesb upon flying the helicopter and more upon other pilot duties. The flignt control sys'em should be designed to allow the pilot to detect or diagnose a failure, disann the failed system, and effect corrective action. This requirement may involve soine for- i of onlinc statusmonitoriag for the variouL control system elements. Design compliance with the current Military Specifications does not preclude th,- possibility of inadvertent flight operationt with one channel of a dualchannel augmý Izstion sy:ni inopcra.v¢. In this type of failurt, the differenct, in flying qialities is small enough to be undetectable by the pilot. Thus, the pilot may enter a flight condition in which fail ,re of the remain;ng SAS channe! cannot be corr cLed within a Zasonable reaction time. For certait critical situations, a need exiets for automatic control activation. For vxainple, electromtchaniLal SAS links should revert automatically to a mechanical lock if hydraulic pressurr is los. .This eliminates the possibility that a sloppy extensible link will create control difficulties while the pilot is attempting to cut off the failed system. Another exampie involves external cargo-handling or -lowing operations, where it may be necesry for the load to release automatically if the applied moments exceed safe levels of controllability. The rerul s of the failure effect analysis (see par. 6-2.3 of this volume and Chapter 3. AMCP 706-203), including any supportintz piloted simulations, should be reviewed and verified 5y flight test. These re.ults then should be incorporated into flight handbooks in the form of wuning notes or flight restrictions for various failure conditions. 6-3.3.2 SAS Faihnur A discussion of SAS failure madu, limitation of authority, and time delay criteria may be found in par. 6-4.4, AMCP 706 201, and in par. 6-2.3 of thib volume. Failurts of rotor-mounted. gyroscopic SAS's
"
j)~ are not discussed. These systems sAll be designed so as to be at leat as reliable as are the rotorcraft primary flight controls.
6.33.3 F'ah-sife Prlmdples Fail-safe desian, redundancy, and self monitoring priniples also are discussed in par. 6-4.3. AMCP 706-201.
L
6-3.3.4 Battle Damage, Vulner.41llty Steps shal be taken to reduce SAS vulnerability in cases where loss of all stability augmentation would abort a mission Duplication or triplicintion of actuators and hydraulic systems is a valid approach. SAS actuators should be designed so that it is possible to lock tthem in a centered position in the event of loss of hydraulic pressure. Duplication or triplication of hydraulic lines does not reduce vulnere bility unless provision is made for automaticall) cutting off the oil supply to sevcred lines. Levers, bell cranks, and pushrods can be made large in size and of lig'st-gage, low-stressed material in order to ir"-uce vuIneainhlitv tn--Imall arms fire. Critical components not readily duplicated should be grouped and protected with armor (see par. 14-3).
j)
6.3.4 COST' 6-3.4.1 Developrmat Coam The cost of developing a new SAS generally is in proportion to the ad vance in the state-oif-the-art repmoented by the dev~lopment program. A conventonal SAS for a conventional airframe can be obtained from off-the-she'f components, whereas a new concept for a rotor-located SAS, or for sensors based upor new teclinoiogy, may require a large expenditure in order to bring it to production status. The new cont&pt must promise a sufficý-nt increas: in cost-effectiverness in future production to rcompensate for the U.gh cost of development. 6-3.4.2 Puodsacton Cost SAS production cost can be reduced by adhering to thc following: 1. Simplicity c-f design 2. Use of integrated and printed circ.uits 3. CommonalitY' ot circuit mod-Ales 4. Extensive use of value engineering principles, Production cost increases may be expected with anl increase in: 1. Number of system components 2. Quality or precision of components 3. Number of nonitandard parts \ 4. Number of parts that can be assembled in\corrft-tly
WrC 706-202 S. Number and interdependency of adjustments to be made in final assembly 6. Number of parts that can be damaged easily in assembly 7. Dogrce of cicanliness required during assembly 8. Unrealistic requirements, or ovcrtzmphasis on snua icpiesc n a. Wcght ri~duction b. Compactness c. Functional complexity d. Reliability e. Maintainability f, Structural integrity. 6-3.4.3 Msllaterancc Cvt A simple, mechanaical SAS composed of infinitelife parts (as in the Bell s".bilizer bar) requires maintenance only in the form of regular inspection and lubrication. In the event of battle damage or other failurc, repairs can be performed by a qualified mechanic. An electrohydraulic SAS, on the other hand, may require the sei vices of an instrument snecialist- an electronic technician, and a qualified helico.pter mochanic. Thus, the ucif-temt cireutit should be devised so as !a indicate exactly which section is defective. Removal and replacement of plug-in modules represent field mainteniance at lowest cost. Added to this cost, however, is the cost of maintaillilg adequate spares. 6-3.5 TECHNICAL DEVELOPMENT PLAN For the dcvcelopmeni of P conventional (e~lcctrol'yd..aulic) SAS, the plan outlinrd in MIL-C-1g244 should be followed. In addition, the airframe and rotor dynamic and aerodynamic propcriics cvcnwluly shoull he inclucied in the initial system analysis (MlL-C-48244) in order to show thec possible existenc of a~rame cross-coupling and the need forr anticross-coupling in the SAS, as well as to show the overall behavici of the SAS/airframc coml~ina' ion. Six degrees of freedom of the airframe, and quasincrnial modes of the roior (inplane as well as flapping motion o,' the bltdes), should be used. The resulting equations art used later in the simulation studies required by MIL-C-18244. The simulaticn not only will 41low the pilot to evalute the system, but also will permit demonstration of the scverul types of failure of the SAS, and will indicate time delays permiscible ;)eibre starting corrective action. In the development of unconventional SAS's, eapecisily those involving modified rotor dynamics, s,-vwal changes from the procedure in MIL-C-18244 ar mommended Unconventional systems require more inisial system synthesis, or concept selection, 6-13
A M M7-2U2 than do conventional systems; and model studies should be undertaken as an aid. The models can range in complexity from simple mock-ups of gyro and linkage arrangements, through dynamically scaled wind tunnel models, to remote controlled flying models. Tne paragraph of MIL-C-18244 dealing with model studies notes that experimental models may take the form of full-scale, engine-driven rotor and SAS assemblies, suitably mounted upon a truck bed for measurement and observation of dynamic behavior under forward-flight conditions. The maximlium possible experience with and knowledge of the system should be gained before the start of testing of a man-carrying flight article. Full-scale wind tunnel tsts, although expensive, .-an be used to test the flight article progressively to conditions beyond the extremes of the projected flight envelope. Further substantiation of the airworthiness of an unconventional SAS and rotor system can be obtained by operating an identical system on a tieiown teat, where a given nurmber of hours is requirc-i for each hour of aciuai flight icsinig. The documentation and data rcqulrcd to establish the satisfactory fulfillment of the technical developmant plan are described in MIL-C-18244, substituting SAS for automatic flight control system (AFCS).
6-4
PILOT EFFORT
The helicopter designer must consider pilot effort, or control system loads, from two points or vie,.. The first concern is the significanice of cntrol feel with redtr by the normally fly nd qualities. phyicald t fiying to ssocfin gard atios. Pilot. ofappliednorcea physical association of applied force and the mapeuve'ing response of the aircraft. iherefore. the cont.ol feel in maneuvers plays an important role in the assessment of handling qualities. Stick positioning also is a fundamental characteristic, be cause it h3lds the helicopter in the selected trim attitude when thu controls arc released. MIL-H-8501 provides for stick position trim and hold by specifying breakout forces and force gradients. The other design consideration is rclaied to the structural integrity of the components. The components shall achieve specified factors of safety when subjected to loads due to pilot and copilot effort, artificial feel devices, power actuators, etc. MIL-S-8699 wefort. covers this aspect of pilot 64.1 CRITERIA FOR POWER CONTROLS Whenever the magnitude and line-.rity of control 6-14
:-•
loads permit, direct mechanical control saill be uwed unless there is a valid requirement for power controls. Direct mechanical control is the simplest and most foolpioof control system. Howcvcr, power-operated systems may be required wbhn the control system environment contains high control forces, feedback of vibratory forces, or mixing of control forces. The control system designer must verify a need for poweroperated systems before adding their cost, weight, and complexity to ,he helicopter design. 6-4.1.1 Control Forces Not all helicopters require power actuatots. For cxarrple, on small, single-lifting-rotor vehicles, a system of weights may be installed in the antitorquerotor controls. Centrifugal force acting upon the weights balances the pitch link loads, and the system is adjusted on the ground to compensate for the control forces in cruise. The pilot t•annot trim the system in flight, and accepts the unbalanced forces in the pedals in hover and flight modac other than cruise. However, this is a small disadvantage in comparison with the aimplicity of mechanical design. Also, it may be feasibie to design a bungc spiring th.. wdil counteract the steady download in collective pitch ' Medium- and heavy-lift helicopters generally require power actuators due to the magnitude of their pitch link loads. Pitch link loads are sensitive to rotor blade design parameters, both aerodynamic and inertial. 6&41.2 Vibration Feedback The control moment of a lifting rotor blade is a steady pitching moment with various alternating harmonic components superimposed. in the nonopoet system,eermoc.Intenn rotating control these components appear as n-per-rcv forces due to the n number of pitch litiks passing over the attochment point where the nonrotating controls support the swashplate. The preoence of these vibrations in sh cyclic p stick generally is intolerable to the pilot. Vibration absorbers can be used to reduce the amplitude of the vibration transmitted by means of a simple mechanical system. 6-4.1.3
Kinematic Effects
The generation of control forces and moments along and about ¶he various axe of the helicopter is accomplished by combinctions of collective and cyclic pitch on the rotor(s), as discdssed in par. 33.3.1.3, AMCP 706-201. The motions of the cyclic stick and thrust lev,;r (and, on sonic helicopters, the motion of the pedals) are transmitted through the swashplate to the rotor(s). In some installations, the
. .
AMCP 706r202 control input
are transmitted to control-mixing
assemblies, vhcrc they arc combined botforc reaching th: swashplate. The degrees of rotor blade angle trol travel in the cockpit, are the dominant consialcration in establishing the mechanical ratios in the
INUUT B
UTPU1 A-6.)
_1 _.
INPUT A L _ 0,1
o0) OUT'PUT (A+B)
(A)INTRIM POSITION
mixing Even if control forces are low and the vibratory components insignificant. there is a €,oss.
talk of forces from one control ax-s to another because of the mixing. It is unlikely that the mixing assemblies, whic" contain components sized for stroke or travel relationships, will produce satisfactory force relationships. MIL-H-8501 sets limits upon control force cross-talk. Fig. 6-4 is a schematic diagram that illustrates the mechanical mixing of mixing coitrol assembly.signals. Fig. 6-5 shows a mechanical
I j.--s OUTPUT
INPUT B I-
o
INPUT A
f
(A-)
4•
OUTPUT (A+B) (B) POSITIVE VA._IE OF INPUT B FROM ýR.I
t
4
Figure 6-5. Mechanical Mixing Assembly
"%
INPUT B
OUTPUT OUT(A40 ) -
INPUT A
6-4.1.4 Control Stiffness At high airspeeds and disk loadin3s, the onset of rotor stall flutter can limit the flight envelope. One of the many parameters to bc considered is the compliance (stiffness) of the contrc! ,y,'-. particularly of the swashplate and. its support. Hence. another justification for power actuators is based upon rotor performance. Fig. 6-6 illustrates the installation of power actuators for tandem helicopters. 6-4.2
OUTPUT (A+B) .0 (C) POSITIVE VALUE OF INPUT ArROM TRIM Figrc 6-4. Control Mixing Scherntik
HANDLING QUALITY SPECIFICATiON
The handling quality requircinants of MIL-H-8501 sshal. be specified in the detail specification if the rotorcraft under design is a pure helicopter. However, if the rotorcraft is a high. performanc viehicle with fixed wings and alternate means of producing 6-15
horizontal thrust, the detail specificat-ons may specify requiremesits from 5oth MIL-H-'8501 and MI4L-F-785. The requirements for contrul feed forces in normal helicopter operations are found in M11-41-8501. The maximum and minimum breakouts and force gradients are defined, alor~g with th-, limit formes. No gradient is specified in thrust, because a collective sick holding system - e.g., adjustable friction om a bpc-
gcnerally is provided. There is no require-
ment for any gradient cxmcpt that it be linear from. trim to limit force. MIL-H..SS0l identifies the maximum control feel forces that arc allowable after a failure in the power boost or powur-opcrated system. The limit force in the failed mode is larger than, but of the same order of magnitude as, the limit load in the normal operating mode. Consequently, if hydraulic boost is requ;.ed fc. normal operation, dual boost probably will be required for the failure mode.
STATIONARY SYSTEM
12ROTATING
SYSTEM
-
~
ROTOR BLADE
I0
POWERED
ACTUATOR
(
STATIONARY SWASHPLATE AUTOMATIC (CYCLIC) TRIM ACTUATOR Figure 6-6. Powered Actuators (Tandem Helicopter) 6-16
POWERED ACTUATOR
4.4.5 u-. 3
IJMAN. Apcr.quii~ofora~ ectve ystm cntrl i a A per~qu~it Wan ff~vcsytemcotro i dcugn dvfiiti~on ofcoiritrcil au~mentation nee~ded as function of tota pilot workload. A force feel system may require no Pilot Control in order to maintain a trimmed flight condition. -
Ceutmi Force Co" "Thcontrol force sytem should provide. 1. Trim position identification tha: will enable the pilot to feel an out-of-trim condition and to fecl and identify trim when returning 2. Hold control in trim when the pilot is flying hands-of 3. An increased force cue to indicate increasing sevcrity of mantuvciring wherever it occurs. An increase in gradient with inciea'ing airspeed is recoinmended. Care should be taken tu avoid force cues introduced to the longitudinal control due to collective inputs. The optimal system would provide a constant relationship between longitudinal stick forces and resuiting aircraft load factor during maneuvers. The control force fc.zl system provides an im6-43.1
mediate and :lignificant cue to the pilot, indicating the
*
woptcr ricpoajonh to control conainand in any flight ct.. Jition. This tightens the loop of pilot control and vehicle response, and enables the pilot to realize optimum control. A lesser performance leaids to use of the feel system only as a trim hold device, and the pilot may prefer to turn it off tinder demanding control situations.
*
6-4.3.2 Dev'dopme,.tal T"~ Moving-bamc flight simulationa can be useful in developing the optimum control feel to suit the heclic pter mission. In the movingt-base simulator. pilots can draw upon past expcrience to identifydesired force ;eel characteristics. Stick force pro-
*
~portional to rates of control displacement, helicopterK angiular rates, and to normal accelerations should be investigated so as to insure the design of an optimum system. As the functions of the artificial feel system are ipcreased, the complexity of the feel unit also incicases. A design requirement for a specified linear gradient in the region of trim and a different linear geadient at greater excursions can result in a feel system with more than one spripg. Furthermore, ifsa rtquirement exibts for' nonlineapr forct versus doflection characteristics, cams or linkages can be employed. Fig. 6-7 is a schematic diagram of an ar'jficial feel system. Flight safety at high speeds can be increased by reducing thc occurrence of high rotor loads associ-
ted with excessive control displacement. Adynamic pressure-scnsitive (q-scnsitivc) control force fedl system produces minimum forces in hover sand maximum force gradients inh.gh-specd flight, wherc the sensitivity is ~rrcatept. This concept is an alternative to Use Of Acontrol ratio changer in the pnmary control !ankagc. The q-feel system car~ be modhanical (with q-bollows), electrical, or tlectroicat -el spoie u hydrauslic, In fixed-wing icat -el spoie u slight'y different purpose. Thc pilot flies the aitplane: by sensing, among other cues, normal acceleration and control itick forces. Response of an aiipJAne IS suhta4h cag nnrm!aclrtinjrui of elevator deflection increase with q. If the artificial stick force per unit of elevator deflection also is made to increas with q.then the relationship of stick force to normal acceleration can be made to approxima.te a constant value of stick force per g, regardtess of flight speed.N Military Specifications useful in the detail design of the artificial feel system include MIL-H-85fl1, MILS-8698. NIIL-F-8785, MIL-F-9490, anid MIL-F18372.
IC
pa.
CAM
C tTi0L r_01
r-
PWJACATF
--
tip
ITRWM coNT"V CBE
Z1AO
(A)TRIM CONDITION CNRLCNEE 7
'±
-.
(B)TRIM CONDITION CONTROL DISPLACED Ftgure 6-7.
CONTROL DISPLACED RMTI
CONTROL DISPLACED FROM TRIM
ArdtWIca FWe sad Trim Schematic
64.A AUTOMATIC CONTROL INTERFACES Inner loop stabilization signals are summed with the pilot's commands through electro/hydromechanic&l actuators in series wi*.h the pilot's controls. It is important that the high-ftequency, small-amplitude stabilization signals do not reach the cyclic stick in the form of formes or deflections. Thus, there is a 6-17
~ -
AMCP 70620 need for a "no-back" (a device to prevent the foed-
data were taken and the environment in which the
back of forces) located upstream of the SAS series actuator. A stick boost also will perform this function. In addison, if the helicopter is to be equipped with an autopilot that introduces signals through actuators that move the cockpit controls in parallel with the pilot, there is a requirement for compatibility amr.ong
new system will perform must be assured, or appropriate adjustment of the projected rates must be made. Another rationale for duplication is based upon failure considerations. A power actuator may provide the required reliability; but if a failure of the ac-
tVe inertia, compliance, and damping of the primary
tuator is catastrophic, a redundant actuator is re-
m.chanical controls and of the parallel actuator.
quircd. Further discussion of this subject is con-
tained in pars. 6-5.2 and 9-2. 6-4.5 VULNERABILInY The close support of grounid operations exposes the U S Army's observation, cargo, utility, and armed helicopters to small arms and automatic weapons fire, The unprotected, singlc-channel flight control system is vulnerable over its entire length. There are a number of weys to reduce this vulnerability. One naethod is to make the components so rugged that they can sustain a hit without losing their struc-
6-5
MECHANISMS
The rotating controls in the main rotor sytem nor-
mally include the rotating swashplate, the pitch links, and the drive scissors. These components are shown in a typical arrangement in Fig. 6-8. Functionally, the rotating swashplate translates along the rotor shaft
and tilts in any plane as dictated by control inputs.
tural integrity. However, this is seldom feasible, especially when space and weight must be controlled rigidly. Cert.,m areas, such as the cockpit, will be proItectec with armor piate in order to safeguard the
The swajhplatc translation and tilt arc transferred to the blade pitch horn through the pitch links and, thereby, control the main ro.or thrust vector. The of the rotating con(a)ihc fix ruoor tMe positionp scissors iu drive 'haft and rotor blades. and cruis rchaivc
crew. The same armor can te used to shield the mechanical controls. However, it may not be feasible to run armor plate all the way to the swashplatc. A redundant cintrol system not only helps Wo solve the vulverability problem but also improves flight safety reliability. To be effective, redundant channels must be separated physically. Consideration must be given to single-channel jams and disconnects, to adequacy of control if the remaining chann-.l goes to half gait, and to the question of whether the configuration should be active-active or active-standby.
(b) provide the load path for the conversion of drive shaft torque into the tangential force required to induce rotational motion in the rotating controls.
64.6 RELIABILITY
the rotor blades complete one revolution, the pitch
6-5.1.1 Design Factors Structurally, the rotating system shall be designed to withstand the alternating (fatigue) flight loads introduccd by rotor blade torsional moments and the maximum loads introduced by severe flight mameuvcra or during ground operations. The fatigue loads are periodic, and alternate primarily on the basis of once-pcr-rotor-tevolution. In other words, each time
The overall reliability of a flight control system depends upon the reliability of the individual components and upon their arrangement, which may be either in series or parallel. If the helicopter system specification prcscribes a minimum acceptable value for flight safety reliability, this value may be so high as to require dual mechanical controls. The detail
link load completes one stress cycle. Therefore, a high-cycle fatigue evaluation is rcquired. The primary loads are discussed in pars. 4-9 and 4-10, AMCP 706-201. and the fatigue evaluation is discussed in par. 4-11, AMCP 706-201. In addition to the primary Pisht loads, special consideration shall be given to socondary loads. Failure
designer first must establish the single success path;
to evaluate secondary loads properly may lead to set-
then, if system reliability is inadequate (a value less than required by the helicopter system specification), he must add redundancy, beginning with the least
vice problems. Among the secondary loads that shall be considered are frictional moments in rods and bearings, and bending moments created by centri-
reliable componn'nts,
The reliability of a component is a function of the
fotgal force. A typical pitch link rod end, with a selfaligning bearing, is shown in Fig. 6-9.
historical mean time between iailures (MTBF) of that
Bearing motions of -6 deg are not uncommon
component. When historical failuic rates arc used, similarity between the tav,ronment under which the
during each rotor revolution. The normal force (pitch link aad) times the coefficicnt of friction produces a
6-18
•
I)
M~CP 706-202ý
ROTOR BLADE-,
LISTATIONARY
SYSTEN.7',
ROTATING SYSTEM
'j<,
CNRLHR
ROTOR SHAFT DRIVE SCISSORS
SWASFIPLATE RING
Figure 6-.
Rotating Controls
6-19
AMCP 706-202 frictional force upon the spherical surface of the
PAPLIEO AAIAL LO•A)
bearing. Rod end motion, in the presence of frictional forces, induces pitch link bending moments. If the
rod end and the bearing arc of different material, diffcrential expansion due to temperature changes will alter the frictional moments. The b-nding stresses that result may be significant and should be evaluated
at the same time as the primary loads. A pitch I:nk bending moment also will result from the centrifugal force acting upon the weight of the pitch link. This inertia force will produce a transverse deflection. Although this deflection may be small, its effect upon pitch link strength shall be evaluated from a beurn-column standpoint (see Fig. 6-10). The ultimate and limit strengths shall be sufficient for the maximum static leads resulting from both
..
DE
`'
IKLT
•'mwc._.Mi4 c I
Of
rowi
-
AXIAL LGAES A
"P Figure 6-10.
Centrifugal Force Deflections
flight and ground operations, including loads during
the blade folding 'f applicable. The sources of these loads are discussed in pars. 4-6. 4-7, and 4-8, AMCP 706-201. Compensation for tolerance buildup in the rotating control system and the rotor blade usually is provided y nit.ch link e..ngth adiustment. Threaded rod ends are common. I he adjustment provision requires close design attention. Positive Iccking features shall be provided in order to prevent any length change after system rigging. Such changes could be induced either by inflight vibrations or during routine maintenance. The pitch links shall include inspection provisions so as to assurc that sufficient thread engagement is present to provide structural integrity. One method is to provide an inspection hole. Fig. 6-11 illustrates a turnbuckle type of adjustment, showing the inspect-on holct, and installed 1,...ire. The J-m nt
JA
NuI
TURNBUCKLE KEYED WASHERS INSPECTION HOLES .
JAN NUT
ap~st.~s.LOCKWIRE
-
-7-
NOTE: ARROWS INDICATE BEARING MCTIONS Figure 6-9. Typical Pitch Link Rod End
6-20
Figure 6-11.
Pitch Link Adjustment Frovislons
-
~
--
-
-..........
- <~-
--
AMCP 706-202 detail design of 'he rotating system.
SWASHPLATV PITCH HOR'4
-Although
LN
/ -
-
-.
-stress
S""certainties I
/
SWASHPLATE
gRand
SFigere 6-12. Relatve Flich Link Rod End Position forces a keyed washer against the turnbuckle, and is attached to the washer with a cotter pin. The nut/washer combination prevents turnbuckle motion that would shorten the rod, and the lockwire prevents any rod extension. The relativa position of the rod ends is maintained by an internal slot arrangenent. Proper relative rod end position is imVortail' in order to assure sod end clearance within the swas'p.ste and pitch horn loss (Fig. 6-12). In addition to the rotation rriative to the stationary controis, relative motion occurs within the rotating controls. The drive scissors has two horizontal pivots and one universal joint to accommodate the vertical and tilting motions of the swashplate. Pitch link rod eiids have self-aligning bearings to accommodate the small angular changes between the swashplate lugs and the pitch horn caused by swashplate motions. Rod-end-to-lu% clearances a/sol be provided in order to prevent contact during these motions. Ale Lrjjjll W. ... dAMi.. 4r :ce , Although wear may reouce frictional moments, the vibratory levels tend to increase as a result of the ;ooseness caused by wear, and bearing replacement becomes necessary. Ease of bearing replacement is a design consideration. Bearing replacement times are catablishcd hy TBO test programs and by Pervice cxperience. 6-5.1.2 Test Remlts As detail design progreses, it becomes pnosible to replace preliminary design estimates with quIntitative information gained during bench and flight testing, Chapters 7, 8, and 9, AMCP 706-203, define the procedures, tests, and demonstrations involved in demonstrating proof of wompliance with the design requirements. Chapter 4, AMCP 706.201, describes the procedures for fatigue-life determination. The "discussionthat follows defines methods for insuring that bench and flight test data are sutfficiently timeily and complete to be used to best advantage in the
6..!.,.1
Belch Tests the fatigue analysis of the rotating controls may be thorough, the effects of the complex concentrations introduced by locking features and threaded connections, along with other unsuch as fietting, preclude an acceptable analytical fatigue strength determination. Therefore, it is essential that bench testing to determine the fatigue strength of components be coordinated properly with other elements of the design prucess, that the fatigue test requirements be based upon representative .- or. at least, conservative - service conditions, The factors discussed in pars. 6-5.1.2.1.1 through 6-5.1.2.1.4 influence the estab!ishment of the test requirements. Component fatigue test requirements are discussed in dezail in par. 7-4, AMCP 706203. 6-5.1.2.1.1 Test Loads Although it is desirable technically to duplicate all flight loads un the bWach, th's is not always an economic or physical possibility. When flight loads will not be the basis for bench test loading, an analytical assessment must be made in order to determine which of the secondary loads is significant. In the ro.ating conrol system rod end. frictional moments are usrally significant whi!e centrifugal forces are insignificant. Steady loads in rotating control system components generally are low in comparison to the alternating loads. Consequently, the load range is thrwh zero.. thus increauin,, the relative motion of components and the possibility of fectting. Test loads should be programmed so as to insure loading through zero. If the moment induced by rod end friction is significant, :1 must be included in the test. This secondary load must be phased properly with the primary load. The effects of end moments may be induced artifically by applying eccentric axial loads. Another method is to use stiff bearings and to induce bearing motion during the test. In ,ither test, it may be neceary to evalute temperature extremes. 6-5-1.2.1.2 imstrnmentatilm The correlation of flight loads to beach test measurements is a primary consideration. Unless the load distribution upon the part under test can be as. certained readily from applied loads, bench test specimens should be instrument4d and calibrated. Where a complex bending moment exists, a component should be instrumented with sufficient bridges to de6-21
AMCP 706-202 ilrrnine that distribution. The location and type of instrumentation sAdl he the same as is employed in the flight load survey. 6-5.112.1- Quantity sad Selection of Sedmens A minimum of six specimens of each component is required for definition of an S-N curve. Where tolerance ;s a significant factor, specimens should be selected from those at the adverse end of the tolerance band. Dimensional tolerances of critical parts generally are tightly controlled; therefore,
-
-
RADIAL BENDING BRIDGES
•'
614
u
special selection on the basis of dimensions usually is not required. However, selectivity on the basis of more highly variable quantities, such as rod end friction, is required. 6-5.1.2.1.4 Interpretathon of Data If all significant secondary loads are accounted for during bench testing, an S.N curve and endurance limit can be established as a functioi, of the primary alternating load. The endurance limit must be based upon a statistical reduction of test data so as to account for scatter. In some cases, a further reduction f........ L-L . ---ly-Ity.i order t count for a secondary effect not included in the original test program. The preparation of an S-N curve from fatigue test data for a limited nunlbcr of specimens is described in par. 4-1l, AMCP 706&201. 6-5.12.2 Flight Tests The characteristics of the alternating loading on fatigue-critical components are detemniried by a flight load survey. A statistically significant data sample should be obtained for each flight condition represen, tative of helicopter usage, i.e., for each condition ,within he mission profile. "e requirements ior a flight load survey are described in detail in par. 8-2, AMCP 706-203.
BENDING
IDGE
A-
TENIN2.
AXIAL TENSION
6-22
Iarmemr
Pitch LUsk
I
--
TANGENTIAL ENDING BRIDGES
Drive Scissors
6-.1.2.21 Required Imstrmenmatiou Pitch link axial load and drive scissors bending mnonent in the p!ane of rotation are the primary loads in the rotating control system, and must be meuued. Secondary loads requiring measurement are pitch link bending and drive scissor radial bending. At least two bending bridges am required in order to determnine the distribution of each of the moments.
Typical instrumentation of a pitch link and a drive scissors is shown in Figs. 6-13 and 6-14. As a rule, each of the pitch link• ts instrumented with a tension gage in order to determine whether or not there are any differ:nces between the loads from the individual blades. 6-5,1.2.2.2 Flight Coadid-t Flight loads shall be obtained for all mission profile conditions at the most adverse altitude(d) and helicopter configuration(s) within the anticipated operating regime (see par. 8-2. AMCP 706-203). Loads in t~ie rotating control system generally are noncritical in unstalied flight. They do, hower, react to the onset of moment stall and, therefore, usually establish the structural envelope for stalled conditions. Consequently, as a minimnm, control loads should be measured at the conditions most conducive to stall. These are: I. Maximum gross weight Most extreme CG 3. Maximum altitude 4. Minimum rpm 5. High load factor. 6-5.2
F! re 6-13.
-
NONROTATING SYSTEM The location of push rod&and cables must be derwmined early in the duign of a helicopter, prior to Ithe selection and location of other large equipment, so that it will not be necemsry to route the control sytem #round this equipment.
I-)
AMCP 706-202
.)
The control runs must be coordinated with each other and with the entire airframe in scaled layouts. Direct. straight-line routing improves the control system response, reduces friction and weight, and incrcass reliability. Other factors that muest be con-
and power control syterns. In addition to ti. requ.'rmareis of the Military Specification and of Chapter 4. AMCP 706-201, the following iist of requirements and general •.dmign practkm are applicable to tic dsign of helioo1 cr
sidered during control system layout are vulnerability
control systems.
to small arms and automatic weapon fire, jamming by foreign objects, rigidity, strcngth, accessibility for inspection and service, and techniques to prevent inco. rect assembly. All pertinent inforriation should be shown clearly or, the smul!-scale layouts in order to' verify the design feasibility. AS various design options are devuloped, the a&ternative configurations should be evaluated by means of a trade-off study. The parameters in the study may include - but are not limited to -- performance, reliability, cost, safety, weight, use of standard parts, logistics, maintainability, and vulnerability. In such a study, weighting factors may be assigned to the various parameters. However, because the weighting factors affect the outcome of the study, t-.. must be assigned judiciously. The results of the trade-off study upon the selected
I. Push-pull rods, bellcranks, and lovers: a. Each bolt, screw, nut, pin, or other fastener whose los could jeopardize the safe aperation of the helicopter shAwl incorporate two beparate locking device. The festener and its locking devices should no, be affecti4l adversely by cnvironmenmal conditions. b. Impedance bolts shall be used where loss of a bolt can cause a catastrophic failure. c. Rod assemblies should be desigrned with only one adjustable end fitting. The adjustable rod end check nut should be ,ock-wired where the rods are subject to vibration or to hiph-frequency load re. versals. d. The natural frequencies of push-pull rods should be checked against the forcirig frequencies of the rotor(s) in order to assure that the system is free
configuration should be cvaluaitd uanicUilly uiuaii a
fromm on.n_.
design review. The purpose of the review is to insure that the seleted configurations and the applicable specifications, mock-up, and test requirements are in accordance with objectives established during the preliminary design, and that program and contract requirements for performance, rtliability, cost, safety, maintainability, standardization, and ease of inspection arc or will be met.
e. The lagest diameter and longest tube cov,sistent with weight and strength consideratiois should be used in order to provide reduced , ulnerabiLty. f. Maximum clearance in the clevisjo;nts of the push-pull tubes sa be provided so as to allew for overtravel when the controls are disconnected. 2. Torque tubes and universals: a. The natural frequencies of torque tubes should be checked against the forcing frequencies of the rotor(s) to assure that the system is free from . n j t u s r b. Universal joints should be ud where misalignment exists between torque tubes. C. A double universal joint assembly may be
6.5.2.1 Pilot's Coetrols to Pov -r Actutoi A comprehensive discussion of design standards and requirements for that portion of the helicopter flight control system between the pilot'r controls and the power actuator is found in MIL-F-9490 and MIL-F-18372. Although there are minor conflicts between these specifications, the helicopter system specification generally will define the extent of their applicability. The loads in the nonrotating control system consist of those loads present in the system at all times, operational loads due to pilot forces, and any flight loads fed back from the rotor blades. Constant loadL include system preloads or rigging loads, and loads due to component weight. The values of the control system loads required for design are given in par. 4-9.8.3, AMCP 706-201. Included arm the pilot effort loads applied at each ._. control input, together with their reaction points, SCriteria are provided for dual control systems, duplicate systems, distribution of loads within a system,
)
used to obtain constant angular velocity, provided
that: (1) The driving yoke of one of the joints is 90 dcg oftiet from the driving yoke of the other. (2) Each joint is operatwd at the seme angle. (3) All shafts are in the same plane. 3. Cables, pulleys, and quadrants: a. Cables and pulleys should be used only when distinct advantages can be shown over a system using push-pull rods. b. Cables tend to twist over each pulley. If the twist from one pulley rides onto another pulley, cable wear will result. Pulleys thus should be spaced far enough apart so that no ,egment of cable ruus over more than one pulley during full travel.
"6-23
A!MP6'Q MM c. Unsupported spans of 150-200 In. hi~ve operated satisfactofily. However, cable idler pulleys in long straight russ minimize friction over fatirloads and grommets. Close spacin6 of cables shsall be a odcd. Cables should not pass within 3.0 in. of sVtrutr, equipment, or other cables. stutr e. The angle be-tween the centerline of the cable and the Plane of the pulley should riot exceed 0.5 dcg. f. To the maximaim extent practicable. cables shoLld o un beas lon,coseas ossbletothe neouldrul axis th og r' airfram e astructure. o th neura axso ifaesrcue g. Friction in a cable systcm should be minimized by: (I) Using a minimum number of pt~ll:ys (2) Using the largest practicable pulley size (3) Using the smallest cable diameter consistent with strength and rigidity requ~ireme'nts (4) Designing for the sma'lest practicable wrap angle consistent with maintL...cc of ~--od cable contact and pulley rotation. It. The effect of chang~es in temperature can be a serious problem in cuntrol ca-ble systems, due to the diferecric betwee trie coclticicnts ol therm..al eiX pansion of the aluminum airframe and the steel control cable. The problem is less severe in pulleyless cable systems because higher rigging loads are permissible. i. Tension regulators may be installed in quadrent and pulley assemblies in oi icr to allow for expansion end contraction of the cables without appreciable variation in rigging load. j. Nylon-coveaed cables can increase cable life by damping hig'i-frequency vibrations. k. Cab!t guards shall be used at pointi of tanigency of the cable to the pulley, 4. Chains. The use a, chains shall be subject to the approval of the procuripi% activity. In spite of the inherent advantages, applications of fly-by-wire tec~iniques to helicopter control have been slow to materialize. The substitution of fly-bywire electrical signtiling systems fcT conventional linkages between cockpit and swashplate bas a number of potential benefits. However. before any *fly-by-wire primary flight control system is accepted for produc-tion, a high le-jel of reliability must be asawied. The advantages of an electrical control system will de~pend upon the type and size of the helicopter in which the system is installed. For example, sonit of *Ithe benefits to be expected in large, heavy-lift heai*copters are- improved flight safety reliability and rtýduced vulnerat ility, higher fidelity of control, and reduced weight. The characteristics and capabilities of
Id.
.h
6-24
a true fly-by-wire system arc cot. .pared
b..vari.
alternste heavy-lift helicopter control systems ili Refs. 23 3nd 24. -.. oaAtswttkSsae Design? Pow reerA taes and gtheSan a.-p ortate p r toDofi thequeiropemnsadtnd flgtwto ythabtweoror thon pofwter halctaoptr and the nontrolansyastemb twee thpoeacuorndheorttngwsplc acdfndb I40adMLF~32 In addition to these requirements, irnportant considerations include fail-safe design, structural cornpliance, and control system dynamnics. Various fail-safe approaches are: 1. Stand-by design. Two equal-strength ioad paths arc provided. The secoadary load path is isolatO until primary failure occurs. The primary lead pr-th is visible durirg helicopter inspections. This approach r- uir,:* that each loted path be designed for infinite fc tj insure thiat the components of the se'ondary will last between overhauls if the failure of the prim~ ahi o eetd 2., Load-sharing desirpt. Tht. component is mr~de up of two or inore sections or laminations that arc joinied miechanically. If one element fails, the rcmaining JAreaets have fuei ci d-crariyirig capability. If the elemen'ts art; bonded, the bc-nding agent shall prevent a crack front propagating across the section. If thc component is not bonded, other antifreatting barriers must be employc-d. A section of the w.omponent consists of the maximum practicable number of laminaticns. 3. Crack-detection dca~gn. Dzsign techniqLuer could includc: pressure drop, oil or fluid leak, electrical detectors, and highly penetrating dye. lnadcquatc, struc'ural stiffness can affect the control systcm ia several ways. Deflections of the airframe can introduce input's to the eontrol system, an effect that is minimized by routing the controls close to the ncutral axis of the airframe. When the support structurc is not stiff, the power sctuator also can deflect under load. If the control system output is sensed by the compliant structure, a limit cycle instabitity can occu- and ultimately may destroy the helicopter if it is allowed to proceed unchecked. The desij'ner shall introduce compensating linkage or sufficient structural stiffness in ordci to assure that the control system is inscrititive to deiflections. The primary flight controls are part of the comnplex servo system that determines the transient and frequency responses of the helicopter. In this system, the mechar'ical controls play a small but significant role. In addition to the mechanical controls, perforniance of the servo system depenads upon: 1. Rotor dynamics and aerodynamics 2. Inherent helicopter stability
*
(
AMCP 706202 3. Power actuator dynamics 4. Automatic flight controls: (a) Stability augmentation system (b) Outer loop etabilization (c) Automatic trim systems. 5. Pilot iNthe feedback loop, Frequently, tht mechanical controls and the power actuator are analyzed together as a subsystem. Characteristics of the mechanical controls that have. an effect upon the responses of the servo system are: I. Inertia and balance of control system componcns 2. Damping at control stick, actuator valve, and/or control surface 3. Friction at control system joints 4. Looseness of control system joints, Friction can cause control system hysteresis, which prevents the control stick from returning to the trim positior once it is displaced. The provision of positive centering requires a preload force larger than the value of the friction force. However, excessive control break.uAsforce around the neutral or trim position is undesirable because it rcsults in a tendency for the pilot to ovcr'ontrol the helicopter. ii
Looseries, thQ result of escemive buildup of
tolerances and wear at b.Warings and joints, causes "backlash iq tht control system. The effe4ts of backlash can range from sloppy and unsatisfactory control characteristics to pilot-induced oscillations. Dynamic analysis conducted with high-apeed digital or analog computers not only identifies required characteristics of the automatic systems, but also identifies design requirements for the mechanical identiem.suh dsin system, such as: rnetic I. Balancing of certain control components, par. t,•ildfy thic cyclic and. Cl*'.,..... 1c.... 2. Stiffening of control elements and backup structure 3. Insta!lation of antibackl2sh springs to eliminate looseness 4. Additional damping at stick, actuator valves, and/or control surfaces 5. Establishment of the allowable upper limit for Scontrol system friction, Major aocidents can result from improper or inadequate maintenance of flight control systems. Specific aesign guidelines for maintainability of conIrol systems include: I. Understand the skill level of the maintenanre personil, their operating environment, and the type of errors they are likely to make 2. Replace routine maintenance with on-condition maintenance accompanied by adequate failure warning .
3. Incorporate physical barriers against incorrect assembly and installation of gcnerally similar parts. The design shall insure that the omission of critical fasteners either is obvious during ground runup or cannot result in catastrophic failure in flight. 4. Realize that the same maintenance crror may be repeated in all paths of a redundant systein 5. Human factors engineering should be applied to design for maintainability to minimize human error. 6-5.3 TRIM SYSTEMS The force trim system is provided in order to allow the pilot to reduce the control force to zero when the helicopter is trimmed along a stabilized flight path. MIL-H-8501 requires that, for all conditions and speeds specified, it shall be possible in steady-state flight to trim steady longitudinal, lateral, and directional control forces to zero. At all trim cond'tions, the controls shai exhibit positive selfcentering characteristics. Stick "jump" when trim is actuated is undesirable. Severas types of control force trim systems are described in the paragraphs that follow. 6-5.3.1
Disconnect Trim
The handling quality requirements can be satisfied by a preloading spring in combination with a magnetic brake. The principal advantage of this method is simplicity. However, the magnituo,; of the spring force and the kinematics o" the system may combine to produce an objectionable kick when the magnetic brake is released. A damper in parallel with the magbrakethe willtrim reduce thisisundesirable characittbrake istic; whenever button held, the magnetic is disengaged from the control linkage, and, therefore, the trim, or force feel, springs also arc diigngaged. In a well-designed system, this can be an advantage, as it simplifies the input of small control displacements such as those required for precise hovering control. 6-5.3.2 Continuous TrIm An al'ernative to the on-off system is a system in which the trim is continuous. Upon activation of the trim switch, the control forces are trimmed slowly to zero. In thit system, the magnetic brakes are replaced by electromechanical actuators. This type of trim can be provided readily in helicopters that are equipped with parallel actuators for outer loop stabilization. However, the two-axis (Chinese hat) 1ectrical trim switch on the cyclic stick grip can activate only longitudinal and lateral trim; the directional trim switch must be located elsewhere. I o avoid the trim switch limitation, it is possible to 6-25
r,"
.
-.
•
S
:r..
AMCP 706-20',t design the trim circuit so that when the pilot depresses the trim button, the trim force for any control axis that is out of 'rim is trimmed to zero force. If more than one control axis is out of trim, all axes would be trimmed simultanrously to zero. A detent arrangement disengages the actuator when the zero spring force has been reached. both the rate at which the actuator operates and the authority, or maximum value, of the feel force provided by the continuou3 trim syhtem are significant in determining the acceptability of the system No specific requirements are giveii by MILH-8501.
runaway series trim causs a change in both the control position and the fVcroe necessary to maintain trimmed flight. The magnitudes depend upon tLc authority of th, trim system. In addition to trimming steady-state control forces to zero, a trim system may be used for trimming bf aerodynamic forces and moments (series trim). Trim at the incidence of the horizontal stabilize! may be used in single-rotor loelicopters, trim of the longitudinal cyclic pitch in taneein-rotor helicopter, and trim of the wing incidence angle in rotorcraft equipped with wings. This aerodynamic trim may be programmed automatically or operated manually
Fig. 6-6 shows the installation of an automatic cyclic 6-5.3.3. Parallel and Series Trim Artificial feel forces may be trimmed to zero with both parallel and series trim. Parallel trim involves the repositioning of the neutral (zero force) point of the unit; thus, a new trim position for the entire control system is created. Both the magnetic brake and the continuous trim systems afc parallel. Fig. 6-7 shows the parallel trim actuator in the schematic of an artificial feel system. Series trim involves the insertion of an exiendabie link in the control system between the feel unit and the power actuator, and produces control surface motion with no stick motion. Helicopters with fixed wings and alternate means
pitch trim actuator 6-6 SYSTEM DEVELOPMENT 6-6.1 GENERAL Design and development of the helicopter flight control system should include several forms of testing. Objectives of this testing are to improve the validity and accuracy of analytical mathematical models, to insure proper consideration of the human pilot as a controllcr, and to perr-it refined develop ment under a full-scale environment. IMPROVEMENT
of producing thrust may require trimming of un-
In general, the mathematical model used for analy-
boosted aerodynamic control surfaces. Elevators, ailerons, and rudders may be trimmed by driving a geared trim tab through electrical trim motors with mechanical override provided. In addition, the pitch control may be trimmed by adjusting the incidence an&:e of the stabilizer and the elevator. Trim is provided in order to balance, or reduce to zero, the steady-state control forces that arise from changes in helicopter configuration and flight conditions. The vehicle is flown normally by the primary flight controls from one flight condition to another; after allowing time for stabilization, it is trimmed to fly hands-off. To use the trim control to change from one flight condition to Another is a misuse of the trim system. Trimming the 'chicle into mineuverc results in loss of the capabilit) to retui n to the normal flight attitude if the controls are released, as weli as in loss of the feel for the particular maneuver being accomplished. Constant use of the trim system when it is not required will lower the system MTBF. Failures may occur at extremes of control travel or in an uncomfortable helicopter attitude. An inoperative trim system can create unusual stick forces. Runaway parallel trim produces abnormal control forces, and
zing the helicopter during the preliminary design phase considers first-order effects or characteristics and incorporates data or approximations based upon prior experience with similar systems, subsystems, or devices. A margin of tolerance, again based upon available experience, is applied to these results prior to their assessment with respect to specification compliance. During detail design, it is necessary to improve the accuracy and validity of the inathcmatical model in order to insure credible and cost-effective. compliance with specifications. In this regard, wind tunnel and hardware bench tests are proper tools for engineering application.
6-26
6-6.2.1 Wind Tunnel Test The aerodynamic forces and moments of the total helicopter and its components parts, together with their derivatives with respect to many of the variables required for stability and control studies (e.g., attitude, control deflection, and rotor thrust), can be obtained in the wind tunnzl. Primary interest should focus upon the static stability derivatives M., N9, and 4; the damping derivatives M., N,, L., and Z,; the control derivatives M6 , L6 , Nv, and Z,; the speed 3tability derivative M.. and the flow field character-
.
17X AMCP 706-202 istics affecting the helicopter. These data should be gathcrld over the full range of helicopter configurations and for the complee flight envelope. Wind tunnel testing also can provide important information on inttraxis cross-coupling effects. Measuring the damping derivatives directly in the wind tunnel generally requires complex proc--dures and techniques. K~owever, the time constants associated with the changing aerodynamic forces and moment" often arc small compared with vehicle rr.iponsc, allowing steady-state wind tunnel results to be used with sufficient accuracy in dynamic analysem, For ex.-nple, wind tunnel !n maurti,,cnts of the hornzontal tail lift characte.ristics and the downwash existing at the fail location can be used to calculate the pitch-damping contribution of that surface, knowing the tail arnt. 6-6.2.2 Hardware 3ench Tests Aa lardwhre components .rf the flight control systdemrbe ome ponentsble , theyshoube s cte ontrol system become available, they should be subjected to laboratoey bench teats in order to describe accurately their performanee haracteristics. Tprse results then may be,pdate used to the odata elT *vo__! m.p~oyegl e irn the mathematical models. The bench testing can define items rich as stiffness, fmquency responi%, thresho!d level, and rate .;raits.
"
6-6.3 GROUND-BASED PILOTED FLIGHT SIMULATION The helicopter flight control system design not only should result in compliance with the minimum requirements of the handling quality specification, but also should maximize th: handling quality potential from the pilot's viewpoint with regard to the mission requirements. Therefore, detail design of Cornpnex fliaht control systems renuires that the human pilol. be inserted into the simulation by using either fixed-basc or moving-base piloted flight simulators. Fixtd-base simulation, however, does not provide the pilot with a realistic environment for his body sensors, and forces him to respond unrealistically cr, sometimes, falsely. Thj correct environment consists of a proper representation of the helicopter equations of motion and a proper pilot environment, including vision, sound, touch or feel, and motion. Use of a high-fidelity, moving-base, piloted flight simulator is a cost-effective approach to the enginetering development of helicopters. The flight simulator is an engineering tool that can provide high confidence in design decision-making regarding new systems. Application of moving-base flight simulstion early in the helicopter definition and developmental cycle identifies pitfalls and potential prob-
S~6-27
Iems, and provides Lmeans of generating good flight vehicle characteristics. During the past few years, heavy emphasis has been placed upon efforts to identify and understand the fundamental technological and phyfiological fac. tort involved in the man/machine interface, par. ticularly with regard to nonhelicopter rotorcraft systems. where ths basic vehicle configuration often is dictated by the pilot's control ability. These efforts already have produced a quantum increase in the knowledge of control system theory and criteria, stability, human factors, handling and flying qualities, end hardware design. Because nonhelicopter rotorcraft are advanced systems, their ability to comply with specifications must be substantiated prior to any significant financial expenditure. Piloted flight simulation provides the means for assaessins, and demonstrating 'the adequacy of the pilot/vehicle system prior to hardware procurement. Pilotcd flight simulation should be employed in the development of any system whet.. the pilot is involved directly in the contro! loop. A partial list of flight control design and development studies that cn used in piloted simulation includes: ~ beSsetblt opitidcdoiiaiz I. Susceptibility to piot-induc-. Osculation (Piwi 2. Analyses of failure mode effects 3. Definition of control harmony requirements 4. Height-velocity capability, with emphasis on human factors 5. Handling qualities in turbulent air 6. Design trade-off studies 7. Scheduling and mixing of flight control component functions 8. Optimization of stability augmentation system configuration ann flight control system forces 9. Weapon delivery suitability 10. Conversion mode characteristics and requirements (for non-helicopter rotorcraft) II. Flight teat supplement (pilot familiarizetion, test planning, test support) 12. Autorotation entry and recovery. Piloted simulation is effective particularly in failure mode studies. All types and combinations of failurs can be presented for evaluation of transient response characteristics, profile of pilot reaction, allowable time delay prior to corrective pilot action, the need for fully automatic protection, and the resultant limitations on mission capability. Initially, such studies are conducted using a fully mathematical rtpresentation of the helicopter. As hardware com. ponents become available, they can be substituted for their corresponding mathematical models, thereby enabling refinement of the previous estimates made for items such as friction and hysteresia.
20706-20
"66. FLIGHT TESTS
.
REFERENCESK
The detail design phase oi'a flight control system generally ex~tends into the eariy stages of helicopter ibgh, testing. Flight testing particularly is warranted for thc detail design of new flight control system concepts, or for novel applications of a given system. Flight testing associated with the detail design effort requires that the flight vontrol system be mechanized so that its characteristics may bc altered or adjusted over a limited range. After establiishing the helicopter hpndling qualities with the flight con-
I.B. B. Blakc, N. Albion, a~id R. C. Radford. Flight Simulation of the CH-46 Helicopter, Proccedings, 25th Annual Ns~tional Forum, American Helicopter Society, May 1969. 2. R. J. H-uston, An Exploratory Investigation of Feactors Affecting the Handling Qualities oj a Ru~dimntr Hingeless Rotor Helicopter, NASA TN D-3418, May 1966. 3. H. L. Kelley, R. J. Pegs, and R. A. Champine, cofgrto.th trlsystem se.oisoia Flying Qualities Factors Currently Limiting enigincer can vary the system configuration irt an attempt to improve vehicle handling qualities. Hisas-~HelicopterNap of the Earth Maneuverability as Identifie by Flight Investigation, NASA TN Dsesament of change is based upon pilot commentary 4931. and upon recorded time histories of the important ye4. Mao P. R. Curry and J. T. Matthews, Jr.. hide and flight cL..ntrol system response parameters. ~h SgetdR~rmnsfrVSO Such eve'-titions should consider the total requiregesdRqimntfrVSTLFyg ments of the system over the entire flight enivelope. Qualities, US Army Av~iation Matericl Specific flight test requirements arc discussed in TR 65-45, June 1965. an 9, MCP706-03.Laboratories, Chapers D. Cooper, K. Hansen, and T. Kaplita, ShigleS nd06-03.5. , ACP Chater Rotor Helicopter Dywnaici Fo~lowort Power 6.6.5 DESIGN REVIEW Failure at HighI Speeds, US Army Aviation Whtn the configuration has been selected aind the Materiel Laboratories, TR 66-30, June 1966. design requirements have been identified, the detail j. Davis, Hi. McCaijcriy. i. Kuutnun. and F. 6. design proceehs from the one-half and lull-size layLeoine, A Study of Tandem-Rotor Helicopter simulouts. Stress and weight analyses are conducted Dy, -amicsFollowing Power failure at High Speed. layouts fuli-size of preparation the taneously with US Army Aviation Materiel Laboraitories, TR and detail, assembly, and installation drawings. 65-72, Nov. 1965. Long-lead items and material are ordered in advance 7. B. Kelley, "Helicopter Stability with Young's of the drawing release. Lifting Rotor", Society of Automotive Engineers knowledge a ecquires components the of FDetailing Journal, December 1945. of materials, processes, and standard parts. The 8. Joseph Stewart 1ll, Thse Helicopter Control Rotor, reader is referred to Chapters 2. 16, and 17 for presented at the 16th Annual Meeting of the guidance. Detail rcquiranievits applicable to the interInstitute of Aeronautical Sciences (lAS), January hythe and system control faces between the flight 1948. draulic system amdescribed inChapter 9. 9..M. Vague, and C. M. Scibel, Helicopter StaA hazard analysis &Wal be performed it%order to bilization and Handling Characteristics improvedetermine the design potential for incurring equipment by Mechanical Means, lAS Report No. 59ment failures or human errors that can cause Gc27, lAS 27th Annual Meeting. New York, NY, dethis details 706-203, cidents. Chapter 3, AMCP January 1959. sign evaluation technique. 10. M. George. E. Kisielowski, and A. A. The critical design review, a formal technical Perlmutter. Dynagymo A Mechanical Stability enthe when conducted review of the detail design, is Augmentation System for Helicopters. fabrifor release for ready are drawinort gineering 710 ac USAB967nia.eor a purpose of thedesign, ise. orp thoueet.Th cation ac 710 eor 19AL67.cnia prcueentt.l acceptarplit of the resignw ise. ction oetrmn Fraundorf, E. and Kisielowski, E. George, M. IL. requiredesign the thttedetail design satisfies Reliability Evaluation of a Mechanical Stability se t forth in nd te dsig sout atifie mons, Augmentation Systemw for Helicopters. prior reviews. USAAVLADS Technical Report 69-17, June Review team mebr ersnigproduct sup1969. maintainabiliport engineering, flight test, reliability, 1. Culver, and D. Walters, Tali Rotor for 12. aerodynamics, stress, factors, ty, safety, human Helicopter, U S Patent No. 3,004.736, Octobe* materials, and prtxcess-en'incering shall evaluate the 1961. design as it pertains to their specialized fields. 6-29
I.
AMCP 706-202 13. G. Sissingh, "Response Characteristics of the Gyio-controlled Lockheed Rotor System", Journail of the American Helicopter Society 12. No. 4. October 1967. 14. B. Blake. J. Clifford. R. Kaczynski, and P.
i•
)
Qualities Requirements Based on Mission-Tusk Perform3nce", Journal of the American Helicopter Society 15, No. 3, July 1970. 20. G. Dausman, D. Gebh,.rd, and L. Goland, Development, Flight Test and EWvauation of a
Sheridan, Recent Advances in Flying Qualities of
Mechanical Stabilizer for Singl'e-rotor Heli-
Tandem Helicopters, Proceedings, 14th Annual National Foium, American Helicopter Society, April 1958. 15. E. Ebsen, D. Ogran, and H. Sotanaki, Three-
copters, 26th Annual Meeting, Institute of the Aeronautitad Sciences, Preprint No. 821, January 1958. 21. H. Daughaday, and H. Mc'ntyre, Suppressionof
AxL FluidirStability Augmentation Sjstem Flight
Transmltted Harmonic Rotor Loadr by Blade
Test Report, USAAMRDL TR 71-34, September 1971.
Pitch Control, Proceedings, 23rd Annual National Forum, American Helicopter Socety,
16. W. Bedhun, Fiiudic Three-Axis Stability Augmentation System for the CH-46 Helicopter. Honeywell Dccument 20725-FR, Contract
Paper No. 129, May 1967. 22. H. Hecht, and L. Kaufman, Reliability Requiremcnisfor HelicopterFlight Controls. 251h Annual
N62269-67-C-0086, January 1969. 17. W. Cresap, "D'evelopment and Tests of Multibladed Semi-rigid Rotor Systems", Journal of the American Helicopter Society 5, No. 2, April 1960. 18. H. Edenborough, and K. Wernecke, Control and Maneuver Requirements for Armed Helicopters, Piuini.ga. 2V0th Aaiown: Natiojin Toru,-m, American Helicopter Society, May 1964. 19. H. Harper, W. Sardanosky, and R. Scharpf, "*Development of VTOL Flying and Handling
Meeting instittte of the Acronautiral Sciences, Preprint No. 682, January 1957. 23. D. S. Jenney, and L. S. Szvstak, Control of Large Crane Helicopters, Proceedings, 26th Annual National Forum, American Helicopter Society, Par r No. 441, June 1970. 24. R. W. Sanford, P. R. Venuti, and D. Wood, ........ r C.,.... S . . e--. Proceedings, 27th Annual National Forum, American Helicopter Society, Paper No. 503, May 1971.
6-29
.,
706-202
CHAPTER 7
ELECTRICAL SUBSYSTEM DESIGN 7-0
LIST OF SYMBOLS
charging current, A charging time, min SV 'RS - volt-amperes, reactive - increase in battery capacity, A-hr A, -
7-I INTRODUCTION 7-1.4 GENERAL The basic determinants of overall electrical system design and layout are the demands of the equipment on the helicopter for electrical power, and the physica! and operational constraints ir posed by the helicopter and its mission(s). The latter will include such aspects as the availability of space in the aircraft, the sr.fety requirements imposed by the system, and the weight penalty imposed by the ch,'sen subsystem. -f....;r..
..-
'
...
,., .;of e!rcaol
s,.,.
characteristics and utilization are determined by MIL-STD-704. In general, however, helicopters require 28 V DC for both normal and emiergncy operation of such items as fuel pumps, flight instruments, panel lighting, most avionic equipment, und electrically-driven weapons; constant frequency 400Hz AC power for some avionic equipment; and often variable-frequency AC for some heating (deicing) and frequency-insePsitive loads. Thus, with suitable conversion compokients, the basic helicopter electrical system can be DC, variable-frequency AC (vf AC), or constant-frequcncy AC (c' AC). New helicopter design trends are toward increased electrical power in general, as well as increased 'amounts of constant-frequency power. Many helicopter system designs provide an input speed to the g.nerator that varies by more than +5%. In addition, a major consideration in electrical system selection is engine starting. If the maximum engine starting torque is less than 90 lb-ft. a 400-A, 28-V DC starter-generator powered by two CA-5, or CA-9 nickel-cadmium batteries will provide the simplest self-contained start system. However, this90 lb-ft limit defines a small engine, and, therefore, a small helicopter. The actual choice of elcctrical systems will depend upon the relative demand for DC, cf AC, or vf AC and a weight analysis of the neceusry components. This system selection generally will include tike dccision on the existence (or not) of an on-board
auxiliary power unit (APU). Above a certain total requirement for cf AC, for cxamplc, u considerable weight saving can be reaiized by using a CSD (constant-specd drive) input to the electrical system instead of an inverter system; or the presence of an APU may provide hydraulic or pneumatic starting and thus decrease overall electrical system weight by eliminating the need for a DI startei-goeeralor. A transformer-rectifier would be used in this case for other DC power needs. The electrical power system on the helicopter may thus be based upon such power sourc= or conversion devices as: I. AC or DC generators driven by: a. The main rotor power transmission system b. Engine accessory drives c. Constant speed drives (speed controlled by hydraulic or mechanical torque convertcr) d. Constant-speed turbines (speed controlled by air or gas turbines)
2. Ifivetciis 3. Transformer-rccufiers
4. Batteries. The selection of the type of system as well as characteristics of the components of the electrical system are discussed further in the paragraphs that follow. 7-1.2 SYSTEM CHARACTERISTICS The type of electrical power source generally will have been selected d'urirng preliminary design (see Chapter 7, AMCP 706-201). During the detail design phase, it is necessary to confirm this selection and to define the distribution and utiliaation systems. Detail design begins with the specifications for the particular helicopter, which typically spell out: I. The design gross weight (i.e., the weight of I.,e primary mission payload plus the empty weight, including mission-essential equipment) 2. The maximum Performance capabilitim of the •rcraft at its design gross weight 3. The specific primary power source(s) and the power conversion methods 4. The specific utilization equipment, which will include lights, displays, communication equipment, avionics, fire control, and additional electrically powered equipment such as hoists. While the set of utilization equipment components is typically defined in the specifications, numerous options may still be exercised. For example, it may be left to the discretion of the designer as to whether ans 7-1
S
fiMCP 705-202-
E
auxiliary power unit is to be included. The trade-off analyses involved in system selection are dis-ussed in Chapter 7. AMCP 706-201. Approximate weights of the various electrical system components or typical weights per unit output are presented with that discussion. Only upon analysis of all secondary power requirements and systen,m will it be possible to determine whether or not the requirements for enginestarting, and other secondary power, would be better served by an APU or by an electrical source, In general, duplicate primary electrical power sources will be required. The electrical utilization load will be split betwetn the sources, being distributed on two busses such that the total load will nri I be more than hAlf the capacity of the total source. Thus, in the event if failure of one source, automatic paurP•.eling will enable the remaining source to supply all the electrical pows! required. In addition, an "euek:tial" bus must be provided, All components vital to the safe operation of the helicopter under night and instrument conditions must be connected to this bus. In the event of the complete failure of the primary source, the emergency source (battery) must provide power to this bus typically for 20. min of operation with a IL'* reserve. suoh a specification for the oatttry powered emergency bus typically will require the insta'lation of a battery charger/analyzer in the system. This unit is daigned to insure that the charge is maintained and monitoted and that the power from the primary source is distributed properly during normal c,#cration between the utilization load and battery charging. To illustrate the level of input detail given to thc designer at the outset, the system specification may delineate the electrical system for a particu*4r helicopter -
i.e., the primary AC power source shall
be two 400-cycle, three-phase 120- to 208-V AC
generators mounted on the accessory gearbox, that primary DC sll be supplied by two transformerrectifiers, and that emergency power shall be provided by a nickel-cadmium battery with st'fficient capacity to supply power for 20 min of fligu.t with a 10% reserve. This battery load includes an inverter to provide essential AC needs. Also specified nar the requirements for lighting, communications, navigation equipment, s and other utilization meat components. Even with such characteristics predetermined, it &mnains for the designer to select components that meet the applicable, specification and to insure that the electrical chara-teristics of the overall system conform to the requirements of MIL-STD-704. Substantiation of the system design shall include a load MILE-716.4. analsisinwithacor~ncewit 7analysis 2rdtncein MIL-E-7016. m
7-2
7-1.3 LOAD ANALYSIS The information and general format rmquired for electrical load analysis are given in MIL-E-7016; however, the requirements may be modified slightly (particularly for automated systems) to fit each. program. For instance, the specification requircs that the form be as indicated, but this can be modified to fit the format for automated equipment. One page can comprise the equipment list, equipment description, parts designation, and electrical ratings, while the next sheets can obtain the acutal calculations, with the pages folded such that the columns match the preceding sheet when unfolded. Th,ýparagraph that pertains to operating times can be revised to matc~i more closely the modern generator overload times; i.e., 5 sec, 5 rain, and continuous instead of 5 sec, 2 min, and 15 min. MIL.E-7016 requires that phase-toground identification be A-N, B-N, C-N, etc., and this can be modified to A. B, C, and D, with D being the neutral leg of a three-phase four (4) wire system. However, an explanaeory note must be included; and, because dclta-conniocted loads arc rare in modern helicopters, a code also can be established for this situation and explained. MIL-E-7016 gives the formulp for powcr factor. Usually, this ihiformiation is obtained from the equ-pment manufacturer or by actual measurement; however, the formula in MIL-E7016 may be used. The formula for determining single-phase and three-phase power factors PF ib; connected watts PF (7-1) (connected watts), + (connected VARS)' where VARS volt-amperes, reactive The time intervals for the analysis can be modi,-*.
4.-e ;, .... U
...
..
V~ .
.-
t
-.
rnfic n 'Ar,in
tenths of a second increments, 5 min i. hundieths of a minute, and 15 min in hundredths of a minute may be used. 7.1.4 LOAD ANALYSIS PREPARATION The load analysis, as defined by MIL-E-7016, can be written, typed, or presented as an automated printout. From the beginning of a new helicopter dcsign, a complete elect.ical load file must be kept for each piece of equipment. Ideally, a printed file card should be made, allowing jpace for the following information: I. Name of equipnent 2. Equipment part nui,,ber 3. Rated voltage (normal operating) pha) Type of voltage (DC, AC, single- or three-
...
tl
AMCP 706-202 5. Ampere per wire
6. Volt-amperes (normai operating)
us
7.Power factor (normal operating)7111 "8.Watts (normal operating) -'
9. Volt-amperes (emergency conditions, if applicable)[ 10. Power factor (emcrgt.ncy conditions, if applicable)
II. Watts (emergency condition, if applicable)
12. Opcrating time 13. Source of the above information
individually and recorded on a file card. For an auto-
mated systcin, a computer can do the necessary calculations as a teparate run, or they can be submitted Ifom the rtabulation print-out. After the lilt is as complict as possible, it then is necessary to assign cE-h component a power bus in _-.-
Sportent
aIn'Ck1'. *
sIyaS.Clil~It*~lll IlII1
OJE PANSFQUALj
MhiU it U *UiI
,
,,151200 -I *<1
I
I-
I
INVERTER
It
_o
NoXI
2
28 V DC
.US
"-
15 200V'
]
AC BUSN.2
INV
Most of the prctcJing data can be obtained from the equipment menoufacturer. In addition, the manufacturer of each piece of equiiment should be required to supply a component load analysis. After manufacturer-provi&:d data are recorded. the remaihing information can be calculated. For a nonautomated sysiem, the calculations can be made
V AC 5 No 1
No 2
N,,
28VDC
NBUS 2
N0
l
Ek-.
-
Lf2
POIN, PYACtLJ
SIAARIth,6ENL•"0A
referer, e item number.
FIgure 7-1. 7-1.5 MANUAL FORMAT When preparing a manual load analysis, the figare examples of MIL-E-7016 can be used. A sample power distribution systtm is shown in Figs. 7-1 and 7-2. The generator moanting and drive data and power source output data examples are selfexplanatory. The AC load equipment and AC power source utilization analysis charts frum MIL-E-7016 can be combined on one chart for each helicopter AC power bus. See Figs. 7-3 and 7-4 for sample AC load analysis charts. The equipment components are to be arranged alphanumerically. The typical transient nnalysis is requ'red only for extreme transient loads, and, with modern generator ratings, may not be required at all The engine starting requirements data must b. shown even though the helicopter is started only on ground power and data does not appear in the "Start and Warm up" columre of the load anslysis chart. The information will be used to determine the ground starting power supply requirements. If s battery is used in the helicopter, a chart must be included that shows a theoretical charging factor versus time of operation. This chart wil! depend upon the type and size of the battery to be used, as well u the design of the battery charger.
Typical DC Power Dlstribudts System
7-1.6 AUTOMATED FORMAT For an automated format, the same information is required. The program may be written so that calculations are done automatically as the load analysis is b,-Ing processed. The information then can be chaM.kL. , and be,$ t.,i, giv,-, ,-',''; . A typical automation flow chart is shown in Fig. 7-5. I. this example, a two-card system is used. The first card contains the following information: I. Bus 2 System 3. Item number 4. Equipment name 5. Part number 6. Volts 7. Power factor 8. Volt amperce. The second card contains: 1. Bus 2. System 3. Item number 4. Number of units 5. Phase assignment 6. Notes. 7.3
AMCP 7OC2O2
V
1115/200 ABU
&4S No.1Vo
b!AC
TRANISFOR ER _lf111
fR RECTIFIE
Z
.2VDC BUS No. 2
BU
H
7F
28 V () BUS No.I
Figure 7-2. Typical AC Power Distribution Systeus Ill UUUiUUII.
S
~ ~
~
~
Uih a9~wi3U wUiUi i~tlWudU
'
,*
-
K~ J1II
-
6
,
_
min. and 15-win averagcs for: 1. Start and warmup 2. Taxi 3. Takeoff and climb 4. Cruise 5. Combat cruise 6. Descent and landing 7. Emergency. The third card shown on thc flow chart is for programming purposes only. With an automated system, changes can be madce readiiy, new printouts requested as needed, and individual bus totals obtained at any time during the design of the helicopter. Wi'L' both the automated and manual systems, each bus must be totaled separately.
7-1. ARYbattery SUM 7-1. ARY5. SUM Included in each load analysis shell be a summary of rvsults, which will include a brief summary of generating, rectifying, transforming, and battery capabilatics, compared with maximum, average, and emergency loads. The summary will include any special, limiting, or marginal operating conditions that may exist. The sumnmary will be brief and conicise, and indicate clearly the helicopter power system true conditions. 7-4
M fl
/1 &
RrýLý
tLI14n!KA IVKa
I
A1iqiU
*
r
mois.Jin
7-2.1 GENERAL In the detail design of the electrical syltem, certain fundamental criteria must be followed in order to insure proper selection of power-generation equipment and motors and their applicability to electromechanical energy conversion requirements. Before selection of the electrical rotating cornponenta, certain decisions are necessary. These decisions, which may be preliminary, will form the bases for trade-offs related to the optimization of the entire electrical systemi. These considerations should include the following as a minimum: 1. AC or DC. system 2. AC systems - constant or variable frequency 3. Applicable power quality requirements, e.g., MIL-STD-704 4. DC systems - engine starting roquirements and capacity restrictions Electrical load analyses, includinS any additional load imposed upon the generator by feeder losses 6. Generators -- characteristics of the prime mover; speed or speed range, torque limits, overhand moment (weight) restrictions, and vibration and shock environments 7. Rotating comnponents - details of the installation, including envelope restrictions (length, diameter. tool clearunces, removal clearances, etc.), te-~-
.4j ..... 9 K
AMCP
1
'CTR0CALR UIR9C[NIS Xt 6 P(R IN
l
1! 0 PANT N. .NI
EOuIPMLNI
IJ
OA
B
VA
ATARS INAVlO
__
1
GuNSYSTEm
6G283
I
A
PNELE.PRESS. ND.
2908-22
2
300
3
OXY OIT INO.
AERNOl64,4001
2
300
III 1
NAy. COMPUYE
8712-2.1
1 30.1) 90 3
VENURI 6118.
29036-32-6
1300
b00
LEAD CCGdPSETi SEATADJUST
ANIASG,3.3 D6.4uM Ibb
1 300 2 0.20
1w9 0 IO 393k
706-202
03. 1016.1
'62
6
LOAD t~
V, S _;2
TOTAt.
43
.8
.6
6.6
2.94 7.9
7..
79
66
3t2
2
79
71
3D.8 308
76.
315 315
79
.123
.'2
IS 91
15
16.
.91 97.6
234 2.16
70.
i
5T
2.44
900 104 104
WATTSVAR
,6T26
,Bt 1
¶O 00 6 .90
W6.0
5"71 W 312 234 53u 412
A SE.1SUPP'LE NIARN DATA
1OPLRAIIN N
SLt)'D
VTTI
OR ANICHOi0
AVý VA-S
AVG WATTS
A132 9
T
I.N 16s
ONDITIONSS
START AND3 *A•m UP
AVUWAIS
1
-. 60T
6
-
AGR
AN'
>
,
U
j
TAXI
t
VGAR
VG10'AI!,
6.
A
,S
AI
AVG VA
0
I OIAL
U10
0
16 )
eu
6
8. 41. 1
7 59 42
472
39 6.9
63122 6.
12 64 11
OP ERA T NG
-19
14
CONE'
3)
.9 16.
1
3
TAKEOFF ANDCtIW.C~S
-,
AVGWATT
V
V V~ANSJ
ILY I
0-616 12
0182
15.8
1-2
1.,
12
6.9
L9
6.9
6.9
12
12
lo
12
6.9
16.8 156 15I
]B 16.6. 16.6 '6 .8 18. 6 16.6.166 1 '6 5A61 69
F3492.6
97.5
6.
b 60
SIX
F418
1697
60.)969 WbW6
597
.
6.00
5m
6
59
6.9
5.9 12
158. 166. 150 16
6.05 8 5
506W
97
bo
b
94
473 £2
47n
,• "i• ii •'•" -•i •1•.. !,• •..•, • •
I
S
2073
25
12
12
1216
.
1B.C 6.616.86
2i
268A1 59
6.9
1016.6 I
6.9 .cs
97 6.97.6. 9).5 W6.6 6.06. 60.6 5 t.i
317 112 2U4 134
2114 234
6
Figmrt 7-3. ExampleImmlo
.)m
47 •
V
Is
'6615
31231
F6.22 M.410 43U 63(1 el0
472
At w
A132 0102
41
C14JIst-COi.A1
Tl VG VARI S
ItI
12
T ION!,
Awds~yh AC Left Hand Maim
,7-5 .
,, •,•
•' • r ' • :•
'••
• " l '. •• -Ml ia•• m i,,•
..
P •.,
• , , •,
. . .
12 EQ3UIPMENT
-
~ ~~ -
~ ~ ~ ~
WAW-U STARTANDC
LOAD0R AN'C~i08
9!
PART_~
OTT. INDICATOR 10 1821PXY.
962 I 64-MM1 ABSIWO
I 2
15jtgY~M5G93 0.j> .6 6.00 3.0 C 1.2 1.2 1.2 5.9 b.9 5.9 115 0.71- 11.1 A 16.8 15.8 1b.8115.8 16.8 Is8
L2 VNUINAS A COMUTER 3.2 NAV. FCUI651
*/S33 8712-23 93-25
II
594 C 197.5 97.5 197.6150.6 10 .0 90.3 16 $16 9 0,91-
2
1115 0GAD- 39
12A 12 8 .PS
Il
101 SEATAO.LýST -
DW 1106185
1
C 630
63
50.
0.4 472 4/.2
5.91
1,2 1.2 1.2 5.9 5.9 15.8 15.8 1&11 15.8 16.8 15.8
-0~ 97.5 97.5 97.5150.6 50.4
6.3 6X063
50.8
9tAI412 4.16.3
2 TOIALS 3 TOTALS
Figure 7-4. Example AC La"s Analysis Format
--
.i-.
The exciter functions so as to supply DC to the min&! blocks and/or connector restrictions, main field winding. T-he poles are on the Ftator and temperaturealtitude environment, availability, and on a disturbed widing in dlots on the r, or, Both characteristics of air or liquid cooling members are laminated, as in the. main _ncrator. S. Adequate systemo growth - designer mrust con1 n C'.9iatuf A3 4"A tvk'irUGA~tU&, UAnd Provides LO% tv ocurcc. a power aetecting ad-r future growth when main field through a bridge of rotating rectifiers. the ALERNTOS) 7-22 C GNEATRS 7-2, AC ENERTORS(ALTRNATRS)In order to achiove a self-sufficient generator ovie not depndent upon any external power source An AC power source capable of insuring a power for excitation power - a permanent-magnet pilot exby specified that than, quality equal to, or better MIL-STD-704 can be achieved best by utilizing the citer completes the lc~tsical portion of the AC g-Derator. This generates either polyphase or singleattributes of the conventional salient-pole. synchronAC power, which is rectified either within the our aterntorphase The heheopter ACý generator is comprised of the generator (stationary rectifiers mounted within the niain generator, an exciter, and, in the mejority of generator case) cr in the voltag#- regulator (Jpar.7-4) to provide DC excitation power to the exciter. Decases, a permanient-magnect pilot exciter sharing a pending upon system requirements, the magnetic common housing and shaft. Modern-day alternators pilot exciter may provide control power, protective are brushloss; 6:.I,, no brushes, slip rings, or commucircuitry, or operational power necessary for proper tatots are employed, functioning of the distribution system. 7-2.2.1 Electrhca Desigc 7.2.2. Mecbmalcal Desig The main generator consistt of a stator and a rotor. AC generators are. housed in either aluminum ur uniare The stator is built of steel laminations that housings. The selaction of housing magnecsiunt the co'itaan and periphery inner the orn formily slotted output w~ndings. Thesw windin. are connected in a material is dictated by weight and/or vibration requirements. Lamination eteel for the magnetic cirnormal three-phase., (our-wire mnrnnrcr, and are discuit iseither a silicon or co~balt alloy. The latter conplacrd so as to minimize distortion of the output struction results in a relatively expenisive generato., lamiiof oonsists field voltage wavefrmn. The iotor or but provides a weight advar~tage o' almost M0over a the and "pol,ýs". form to as wo lunched nations system using silicon st*.tl punching.. nuniber ef poles and the rpm fix the output froPrac~ical generator speeds range from 6000 to 12,quency. The main field winding is wound on the rotor rpm for 400-Hz output. Variblefrequency 000 the Je prcivi poles @ad is excited with DC. This in ratings to 120 IrVA art paT~ical to machines sufliciunt provide to necessary mafutiomotive force spews of 20,000 rpm. Applications of 6W0 to 12,000 lUMe of force (flux) In the magnet circuit of the main rp'q require grauz~-lubricated ball bearings with gaweator for adequate, all-load-condition, voltag bearing lives of MW0 hr in an aveane helicopter en2W.1iuron, 7-6
-
Vil
Vl
__
_
_
_
_-_.__
7-2.23
DESIGN DATA EQUIPMEN CHARACTERI_,_os TIS-
CARD LOAD ., ,1DSORT
IPT TAPE
were not considered. In the case of blast-coo!cd generators, inlet airflow and pressure-altitude characteristics of the separate forced-air supply must be defined adequately in order tis.al'ovi proper use of a gcnerator cooled in this manner. Cooling-air temperature versus altitude and ambient tempcrature data are vital aspects of an adequate cooling specif.cation as discussed in MIL-G-6099. Contaminant
EDIT LISTING
--.
ERRORS DETECTED
Cooling
Cooling, a primary requirement in an AC generator specification, may be accomplished by air or by liquid. Both air-coaled and oil-cooled generators are for helicopter applications. For air-coolcd generators, either self-cooling or blast cooling may be employed. In the case of selfcooling, an integral fan is located on the rotor of the machine, and diffusers or baffling are employed to direct the air over the hot internai surfaces. If the in. stallation is such that ducting is provided to the fan inlet, the pressurz-flow characteristics of this ducting must be considered so that all through-generator airflow requirements are met. A self-cooled generator that exhibits good performance in the laboratory may burn up on the airframe because duct restrictions
EUIO 1 0ORMATIONemploked W
EDIT INPUT
706-202
_olingAMCP
_
lprotectiorn of cooling air is required.
Oil-cooled generators fall into two categories, conduction-cooled and spray-coaled. Oil-cooled with inlet oil tempe,-atures generators are practical from -65' to 330 0 F.
YES CANCEL RUN
"Inthe conduction-cooled generator, oil is circulated through closed passages in the housing and rotor shaft. Cooling is obtained by conduction of heat to the oil from the hot windings. The bearings use the oil for lubrication as well as for cooling, and rotating seals are required. The weight of the
Y
ic n- .-. . .
h- - ,-
that of the air-cooled generator.
TAPE
LOAD ANALSSI REPOR
Figure 7-5. Typical Autoration Flow Charl vironm-nt. Oil lubrication of bearings generally is necessary tor generator speeds in excess of 15,000 rpm. Bearing life in excess of 10,000 hr has br-n
~) achieved in helicopter generators using oil I
i-
cation. A typical oil-lubricated bearing AC gene Ator is shown in Fig. 7-6.
In the spray-cooled generator, the cooling oil, in effect, is sprayed directly on the windings. This rest-its in an improvement in heat transfer, along with a weight reduction of approximately 15% compared to the air-cooled or cgnduction-oil-cooled generator. To date, all spray-oil-cooled generators have been applied to 400-Hz sy:tems, and operate at 12,000 rpm. For comparison assume a 90-kVA rating; a modern air-cooled generator using magnesium housing and cobalt alloys weighs approximately 90 lb. A spray-cooled generatnr with the same rating weighs 55 lb. A generator weight of approximately 0.5 lb per kVA is aciicvable with apray cooling. If
spray cooling is used, it is necessary to zcavenge the generator cav:ay, i.e.. to remove excess oil fesulting from spraying of the windings. 7-7
TTRýT
AM-'P
Figure 7-6. Typical AC Generator with Oil-lubricated Berrings 7-2.2A4
I7.
App~lcation Checklist
The criteria that follow should be considcred in the dtaian of all AC generators. This is a minimum list of characteristics that must be defined and adapted to a given application: I. Rating: a. kVA at required power factors b. Voltage at terminal& c. Phases d. Frequency 2. Speed, rpm 3. Maximum weight 4. Envelope, diamiemF, and length S. Mounting detals 6. Cooling requirements
\2.
Applicable Military Spocificatkicas regarding
I' . Waveform 12. Performance requirements under unbalanced
load conditions. 7-2.2.5 Variable-frequency AC Generators Variable-frequency generators are practical in ratings to 120 kVA at speeds to 20,000 rpm. The previous discussions relaive to mechaniacl design, cooling, and application checklist generally are applicable also to the variable- frequency g--ncrator. There are twe significant performancne characteristics peculiar to the variablc-frcquency generator that should be considered prior to its application: 1. Voltage transient performance at high speed Voltage regulation problems over a wide speed
range.
generator and/or system performance, eleciromagneti; interfcience, vibration, etc. 8, Minimum africiency 9.Overloads and time at each overload
Voltage transient performanice at high speed for a wide-speed-range generator (e.g., 1.S: 1)can result in severe system problems. This is because the maximeum voltage attainable from the gcnerator £t
circuit
tainable at the low speed. Upon application of load,
10. Short-circ, tit cturrent capacity
7-8
and time at short-
the high speed isthe speed range times the voltage at-
AMCP 706-202 severe d~ps in system voltage could be experienced. Wihregard tovoltage regulation, whnaspeed range approaches or exceeds approximately 2.5:1, high-speed instability c~n result. Because the regulator is called upon to adjust from an overload at low sped t noloadoWading power factor loads at high orde
of 5:1.portnce.Theweight
7-43 SrARmF/GENERATORS, DC CKNERATORS AND STARTERS Stawe-of-thc art DC systanms for helicopter applications are designed for operation at a nominal 28 V. with power quality defined in accordance with the requirsmnuas of MIL-STD-704. For the majority of applications, advantage is taken of the volumetric efficiency and lightweight properties of the DC starter/generator. Nevertheless, there exiit many applications that, for various reasons, eniploy both, a DCgenerator and DC. starter: The construction of the DC startcr/generator i, '
~comprised of a rotating artnature and stationary
field.d The armature is constructed of Aistack of steel a"laminations uniformly slotted on the outer periphery; the power windingV art connected to the coinmutator and are placed into the slots. The stationary field consists of laminated main poles, interpoles, and a solid steel field ring to which the poles arc attached on tlhe Imner peziphery,
LOn
the main poles are wound the main field win-
dinga, connected either in parallel (shunt) or in series with the armature winding. The interpole coils are conn, -led in series with the armature. A fourth winding - distributed in slots, placed in the pole faces, and connected in series to tht armature - serves to support the main field and to overcomec the do. magnetizing effects of an armature reaction to the magnitic field set up by load currents flowing in the *armature winding. This winding is termed the coinpensatieg winding, and generally is employed with gencertor ratings of 200 A or more. The starter/generator nornuilly require a fifth winding consisting of a single turn in stries with the armature and wound on the main, poles. This winding, during starter operation, aids in increaing the torque output per limpere of input current. Some manufacturers leave thii, winding connected during geerator operation an a differential compound windigthat aids in the tagulation over fte load and snpeedi range. The discussion of the mechianical construction of
the AC generator is applicable generally to the DC starter/genierator as.Ioretochvethe lightest possible weight without sacrificing mechanical integrity, both aluminum t.nd magnesium are used for housing materials, the choice being related directly to th:ehria nvironinefltrequirements. advantage of a DC starter/ gnrtremp3oying this high-permeability material, cmaetoaunit employing silicon stiedl punchings, isof the order of 20*. The starter/generator normally operates from stand-still to speeds of 6000 rpm in irelation to starter mode operating speeds. For genefator operation, it ordinarily '-overs a speed range of approximately 2:1. with 3000 ryi. i the usual minimum speed. and seldom is applied whta- maximum speeds exceed 12,000 rpm. State-of-the-art DC istarter/generators generally employ greasL-lubricated ball bearings. For hclicopter usage, thz bearing life generally falls between 1000 and 3000 hr. The brushes that ride on the commutator and cn duct the current from the pwe surc - in the case of the starter - and to the load - in the case of the generator - are made normally of carbon and copper. Because of altitude requiriements, the brushes are treated with u compound (such as molybdenum disulfide) in order to provide the necessry filming characteristics under c jnditions of low oxygen and moisture. For starter/genecrators, brush life is limited to 5M00D10 hr, depending upon the severity of thec
start. For those applications requiring Senerrtor
operation only, brush lives of up to 2000 hr arc po. sihe.. Air cooling of the DC starter/generator and generator is standard prazf~ice. Contan~i~nn" pro. tection of this cooling air ic required. This cooling may be accomplished by integral fan (self-cooling), by blatt cooling, or by a combination of the two. Pre.cautions are necessary in order to define the cooling conditions adequately. Becmuse of the problem of providing adequate heat transfer from the-brushes and commutator to the oil, oil cooling seldom is em. Pioyod. Following is a checklist of minimum in'.ormatic noo.esaary in order to dcfine ad1equately &stafter/ generator for a given application: 1. Engine type and manufactukcr 2. Intended installation 3. Envelope requirements (diameter and length) 4. Maximum allowable weight 5. Maximum alIawabke overhand moment 6. Engine mounting details
-
7-9
AMCP 706-202 23. Percentage of maximum generator output used
7. Applicrble specifications, if any
at engine cruising speed
8, Type of cooling - blest, self, or other a. If blast cooling, prc.surc available b. Temperature of air 9. Ambient temperature range 10. Altitude requirementsi II. Direction of rotation facing engine pad 12. Engine to starter/gener.dtor pad gear ratio 13. Power sapply for starting: a. Battery, type and voltage b. Giound power unit, type and .altage 14. Engine torque versus speed curves for standard conditions and -65 0 F (or the lowest applicable temperature), plus a notation of whether or not thtsc curves include accessories and gearing IS. Engine light-off speed if not shown on curves 16. Starter cutoff speed if not shown on curves 17. Maximum allowable time to light-off speed
24. Voltage regulator type and applicable specification. For engine starting, either a ground power supply or aircraft battery is used. Ground power supplies generally are of the constant-current type, and provide the best power source available for engine starting. In the majority of helicopter applications, where starter/generators are used, aircraft batteries are employed for starting. The batteries are rated 24 V and are either silver-zinc cr nickel-cadmium (par. 7-3). If multiple batteries are used, they may be connected parallel or in series to provide the desired starting characteristics, but consideration must be given to the applied torque vs generator shaft shear section and the engine gearing limitations. A typical startcr/gcnerator used in helicopter ap-
18. Maximum allowable time to cutoff speed 19. Starter/generator pad rpm at engine idle
plications is shown in Fig. 7-7.
20. Starter/generator pad rpm at minimum cruising
7-2.3.2
DC Generators
speed 21. Starter/generator pad rpm at maximum engine
The helicopter DC generator is identical electrically and mechanically to the DC starter/generator,
speed 22. Required generator output and vo!ftge under all speeds in range of regulation
with the -xception that, generally, no series turn is ciployAd ou thc main ficid winding. The preceding discussion relative to mechanical
I
V
0K
,I
3 SCALE - IN.
,•.
4
DC Starter/Generator
1Fitre 7-7.
7-10
1••'•
:' ••r''•
=
•••'•
:
'•
r
:,•
•
q
,"
,,
• • .
.
.
.
.
.•
K.
AMCP 706-202 design, speed rang, beatings, hrushts and cooling also are applicable here. It is suggested that reference be made to MIL-G-6162 as a guide for specification preparation. A typical blast-cooled DC generato., is shown in Fig. 7-8. 7-.3.3 DC Starters With the advent of the DC starter/generator, the DC starter now has limited application. In the past, the majority of applications were found on heli. copiers employing reciprocating engines, and, therefore, these starters were designed for crankink sCr. vice rathcr than for the type of starting service required by the turbine engine. For cranking. the starter operates at a nearly fixed speed until engine ligh,-off. Present-day starters are designed for starting the turbine engine, and the starter operates over a speed range of 0-20,000 rpm. The DC starter/motor is quite similar, electromechanically, to both the DC generator and the DC starter/generator. Ordinarily, often three windings are used, as opposed to the five windings often
empioyed in gcncriurb. Th, 5tt v ii, ai ",IioIIof't"'
') -
-
~,
lease rate of storage batteries is a matter of common practical knowledge. The engine.startilSg cApability of a fully chariged battery in a 70°F environ.ntnt is reduced to 27% after the battery has been stabilized in a commonly experienced -25°F environment. Experience has shown that the probability of a successful 0*F start capability decreases rapidly as the level of the battery charge decreases. Gas turbine starting problems are not restricted solely to the low end of the temperature spectrum. A less obvious, but very real, problem is encountered at the high end of tha temperature range. The hot start, in which the temperature within the turbine exceeds the safe operating rantije for turbine materials, is encountered all too frequently in high ambient temperatures during the starting sequence of gas turbinepowered helicopters. The hot start can be caused by a number of conditions. A battery that is not charged sufficiently and improper fuel control adjustments are two conditions that can lead to critical hot-start problems. Another cause of hot starts is the timelapse needed to accelerate the engine from light-off to ille speed. The net result of hot starts, whatever the cause, is premature engine failurc. Evmi thiigh a a1', '
main field with the armature and interpole winding provides the highest torque per ampere of current, and lends itself to use with battery power supplies, As in the starter/generator and generator, starter/motors employ grease-lubricated ball bearings. Brushes usually are .of the low-contact drop type, and contain a high percentage of metal so as to minimize voltage drop at the high currents required by the starting cycle. Because of the short duty cycle requirements, an integral fan provides required cooling, Thc DC starter/motor serves no function after starting the engine; therefore, design provisions to disconnect it from the engine accessory drive must be made. This usually isaccomplished by a mechanica!ly or electrically operated jaw engaging and disengaging mechanism. A turbine engine starter is shown in Fig. 7-9.
start might not result in an imanediate and catastrophic engine failure, it will shorten the times between engine overhauls, thereby increasing helicopter operating costs. The characteristics of electrical starting systems (and starter/generator systems) for turbine engines generally are such that maximum torque is delivered from the starter/motor to the turbine upon the initiation of the starting sequence. Starter/motor output torque decreases approximately as a straight-line function at increased turbine speeds, with the starter/mr3tor torque reaching a very low value at starter/motor cutoff. The starter/motor is designed to have a stall torque capability exceeding the torque limit of !he engine accessory drive system. Adding a boaot feature to the electrical starting system provides an additional supply of starting
7-22.4 Boost Starting System Modern military practices dictate the use. of turbine-powered helicopters for a variety of reasons, one of which is their ability to use completely unprepared terrain for takeoff and landing. Therefore, deployment of helicopters away from airfield support functions is common. Deployment in hcO or cold temperature environments having no APU or ground power facilities requires t.Oat batteries not lose their gas turbine engine-start (apabilities. The effect of cold temperature on the energy re-
c-ptually, the boost power is obtained from a combination of gas generator and gas motor, and by coupling the gas motor with the electrical starter/ motor. This boost feature is a supplement to, and in no way is intended as an alternative for, the electrical starting system. Even though the added feature is used only occasionally, the components added to the electrical starter, or starter/generator system, nevertheless must be taken into account, and their effect must be established upon such elements as overhand moment, vibration characteristics, operation of the system in a generating mode, and the overall physical
energy for extraordinary starting situations. Con-
7-Il
Figure 7-.
Biast-aemled DC Guemmfor
Figure 7-9. DC Starter Motor With Solemold-opersted Switch 7-12
CA.-
Flgure 7-10. Prototype Cartridge-boosted EPeiurkal Starter System constraints of specific engine installation. Prototype *hardware is shown in Fig. 7-10. MOORSvironmental 7-2. ELETRICL 7-2.
.-
7) ..
ELETRICL MOORSuniversally
Electrical motors for use in helicopter el.ectrical systems may be either AC or DC, depending upon the primary power source sekocted. AC motors in use are almost universally of the squirrel-cage induction type. They may be three-phase or single-phase, wvith the number of phase to be used being governed prin. cipally by motor size and by the characteristic* of typical load requiremenfts. The squirrel-cae induction motor compfnss a laminated statoi - identical in configuration to that employed in the AC generator - and a'laminted rotor. The rotor has slots on the outer peri~hery containing either copper or cas-. aluminum bars - shortcircuited on both ends - with end rings of the same material so as to form, ultimately, what resemble3 a cage. The input winding is contained ini the stator slots and maty be wound threve-phaze or two-
'squirrel
\ phase.
Housing materials maylbe either aluminum or magnesium, depending ~upon weight and enconsideratimis, and bearings are almost of tho groaeiawbricated ball variety. For the specification of ti~c AC motor, the best guide is contained in MIL-NI-799, The factors involved in construction of the DC generator generally apply also to the DC motor, with the ex~ception that seldom, if ever, arc pole face compensating windings employed. Depending upon tbhe load, the DC motors may be used in series, shunt, or compcund winding configuraiions. For very small motors, pcrrnanent-magnet fields are employed in place of the shunt or series field windings. MIL-M8609 is recommended as a guide for specifiution preparation. In general, the electrical motor may be termed a torque device, inasmuch as a certain volume of iron and copper is required to produce a given torque. The sizz ot the motor is dictated by the torque requirements. For helicopter applications, where size and operate weight are at a premium, it is normal to 7-13
S400
motors at high spoeds in order to obtain a high power output (power equals the product of torque and speed) pet unit volume. This, fundamentally, is why the majority of aircraft AC power systems operate on rather than W0Ha power usually em40- Hit forahr tndhsnrial thj 60nHzcowerial usea. Hmplyayd for industrial and commercial uses. Highspeed operation poses certain problems in bearing and/or brush life, which, when couplea with the fact that the airframe itself has a relatively short life, resuit in overhaul rates measured in hundreds of hours, rather than in the 10-20-yzar life rates considered normal in industrial and commercial ground applications. High-speed, high-power-per-unit-volume operalion points up the design parameter that usually determines the size of the helicopter electrical motor. Increased losses follow increased power output, and the size of the motor must be adequate to dissipate these losses without exceeding motor material .temperature limits. It follows that the availability of effective cooling directly affects motor size. High altitudes, with low air density, will decrease the cooling available from a fail. High ambient temperatures rethe he.. urm,':fer and add to the mofOr tot temperature. Also, there is a significant amouat of beat generated in tlýc rotating member of the motor during acceleration from standstill; and if repeated starts are made, motor temperatures rise 5ignificantly. Another significant item is the effect of voltage and/or frequency variation on motor life and size. These items must be accounted for in the design so that, under the worst conditions, the motor continues to produce the required speed at the spv Tfied torque. This means that, for all other conditions, the motor will be operating at high-speed, high-power output
7-2.5.1 AC to DC Convetters Devices for converting AC into MCare of two basic types: rotary and static. Rotary systems may be either AC-driven motor-generators or synchronous econverters. The latter are essentially DC generators in which slip rings have been connected to the armature winding by equidistant taps. The synchronous converter, in effect, combine the functions of an AC drive motor input with those of a DC-generator output, although with less flexibility in voltage and power-factor control than the motor-generator combination. Converters typically are cheaper, more efficient, and more compact than corresponding motor-gcnerat-'rs. It is important, however, that they operate as near to unity power-factor as possible since their "rating", i.e., relative output, decreases rapidly wiih decrease in power-factor. Table 7-1 displays this relative output relationship for various power factors and number of phases. Converters also must be synchronized with the input AC supply. The relationships between the AC and DC voltages and currents are functions both of the power-factor and the number of phases (hence, number of Elip rings). Converters may bc singic-phnasc, il whiuia nb
there are two slip-rings und two slip-ring taps per pole pair; three phase, in which there are three sliprings and three taps per pole-pair; and so on. However, because of the sensitivity of the output to the total number of phases, converters usually arc operated with six phases. With a sine-wave input voltage, the DC voltage is the peak of the diametrical AC voltage, the latter being the voltage between any two diametrically opposed taps.
At unity power factor and a typical 95% efficiency, the DC and AC currents are equal with three slip of input power, the larger the motor that is required. rings; with six slip rings, the DC current is twice the The electrical motor also is Atorque device in the AC current. sense that andimposed power by input, etc., This are direct results of thespeed conversion firom units torque AC to(TRU). DC is accomplished the load. leads byStatic transformer-rectifier The rectifier to use of motor speed torque curves. These curves concept essentially acts to block conduction during arue,smplyothr sequilibrium opetiorquecu . these cone-half of the reversing alternating current. Thus, are current is unidirectional, but only during one-half of motor. The a peed, current, power input, etc, the sine-wave cycle. B3 combinift, two units and a those values that occur when the motor operates ata ii center tap from a transformer, rectification throughgiven torque. out the full sine-wave can be achieved. A smoothing inductance connected in series with the load in such a device effectively smooths out the wave peaks to a relatively small pulsating rippt e. 7-2.5 ELECTRICAL SYSTEM CONVERSION In practice, most rectifiers c.mploy bridge circuits The typical helicopter requires both AC and DC to achieve such ful-wave rectification; however, the power. Therefore, the ability to volnvert one to the active components of which may be diodes or other as needed also is required. Various types of dethyristors (silicone-controlled rectifiers). The output vices are available by which this conversion can be rulti are a series of square current pulses, which effected, together produce a continuous current output. Ripwin, high•" e.
7-14
hg., -'-
.
9|a
AMCP 7
--
-_
_
_
-202 _
TABLE 7-1. OUTPUTS OF CONVERTERS RELATIVE TO CONTINUOUS-CURRENT GENERATOR CONTINUOUS-
SINGLE-PHASE
THRLE-PHASE
FOUR-PHASE
SIX-PHASE
CURRENT GENERATOR
CONVERTER
CONVERILR
CONVERTER
CONVERTER
100
85
132
161
194
95.5
100
78
120
145
170
90
100
74
109
128
145
POWIR-FACTOR, %
100
ole voltages in the DC output can be avoided by the use of polyphase input supplies.
7-3 BATTERIES In general, battery selection is based upon battery
MIL-STD-704 specifies that utilization equipment requiring ar. AC input of 500 volt-amperes (VA) or more and a 28 V DC output of 5 A or less shall use static conversion, unless it is designed sp-cifically for use with DC generators.
characteristics, electrical characteristics of generators and associated controls, utilization loads, and cartain assumptions in unrcraft operations.
DC to AC Converters Devices for converting DC into AC also fails into
Nickel-cadmium and silver-zinc storage batteries preserntly :r used in aircraft eiectricai sybicvis.a.-
two fundamental classes: rotary and static. The rotary class are basically the same types of devices used in AC-DC conversion - they may be either motor-generator sets or synchronous con¶verters; the synchronous converter having the capability of operating on DC and converting into AC. In this condition they are said to be operating inverted and therefore are known commonly as inverters. In general, what has been said regarding motor-generators versus synchronous converters remains true here the inverteirs are more cmlqcrat, cheaer, and more compact than comparable motor-generator sets. In this DC-AC mode, however, some suitable electrical or mechanical speed control must be used since converters tend to "run away" in this condition. (The highly inductive load weakens the field through armature reaction and allows the speed to increase.) For most ordinary, i.e., relatively low-load, appli. cations the most common DC-AC conversion device is the static inverter. This is a circuit that alternately connects the output lines to opposite side of the DC supply typically via the use of such solid-state components as thyristors. Such semiconductor inverters, using both thyristors and transistors in three-phase, full-wave bridge circuits, are in conventional aircraft inverter in rotaryupon and static use. specific The choice castbetween must depend such parameters Sany
electrochemical system has particular service characteristics. Nickel-cadmium exhibits excellent cycle life and output over a wide range of discharge rates, and is preferred to other systems for starting turbine engines. Silver-zinc gives the highest electrical output per unit weight and volume, but is the most expensive of batteries and has the shortest cycle life. Lead-acid is the oldest of the systems, and is used in helicopter design not requiring main engine starts. It is well-adapted to most conventional electricai circults requiring mcierate discharge rates. lypical comparative characteristics are listed in Table 7-2.
S2,5.'.
. as the available input sources, the required output - chnasacteristics o; the system, and the relative costs and weights. •
•
°"'•
7-3.1 BATTERY CHARACTERISTICS
GENERATOR CONTROL BATTERY CHARGING Engine-driven generators are subject to variations in speed in the approximate ratio of 3:1. Thus, a voltage regulator must be provided in order to maintain constant voltage at high engine speeds. The generator is dropped off the bus by the reverse current relay at all lower engine speeds, and the battery must assume all utilization loads at engine speeds below the cutoff value. Various generator and voltage regulctor combinations have been designed that maintain svrctem voltage down to idling speed. in manner limited, currentwill If thethegenerator dependtheupon the utiwhicih battery isischarged lization load current during a specific time interval. and upon the state o :barge of the battery in that period. 7-3.2
-7-15
TABLE 7-2.
U
"
TYPICAL CHARACTERICS OF 24 V, 34 AH BATERY SYSTEMS .. E IG H T.
'T ' P• EW
- hr / b W
I
~~~RETE•NT ION C U RREN• T
10-13
50 x CAPACITY
9
7
93%
19..;
78
10-13
20 x CAPACITY
5
3
90%
12
1.9
34
2530
6 x CAPACITY
17
16
88%
17
1.4
NICKEL-CADMIUM
75
LEAD-ACID SILVER-ZINC
When the gtnerator output current does not reach
"thecurrent-regulator limit, the battery will charje at constant voltage. The charging current will be determined by the state of charge of the battery during the time .intervals considered. When the generator output current tends to exceed the current-regulation limit, the regulator automatically reduces the " generator voltage so as to lim;t the current to the these current conditions, value.atUnder regulated equalthe to battery thc oafa constant will be cha.rae,• ference between the regulated current and the utilization-load current. The increase in battery stateof-charge 6, in this ".ase can be computed by the followinm formula: A-hrSuch A-hr
(7-2)
where Ic c
- charging current, A - charging time, min -
.1,2.
j
overcharging at constant voltage can result in a condition called *'thermal runaway". This is an un: controllable rise in battery temperature that ultimatcly will destroy the battery. As the temperature increases, the effective internal resistance decreases, permitting ever-higher currents to be drawn from the constant-voltdge source. This in turn decreases the resistance still further, in an ever-increasing spiral. the battery the over-all In general, cha-o"n-- ofSOMA" durinscondition should be monritor,,d lations incorporate control systems (battery conditioner/analyzers) which monitor tewiiperature and state-of-charge, constantly analyze the geneial battery status, and cut the charging proess on and off ar conditioas dictate. systems typically us,; other approaches to battery-charging, some of which are itemized in Table 7-3, along with tweir rincipol operational characteristics. 7-3.3
0.8 - charging efficiency including battery When total system loading charging - exceecs the continuous rating of the generator, a current regulator limits the output current to a safe value. This means that the battery m ust. supply the difference between the utilization load current and the maximum generator current. In any case, when the system is operating, the batte~y is always either charging or discharging at some rate determined by the demands of the overall utilization system and the output of the generator. "The actual conditions of the charging state at any time are controlled by the generator voltage and current regulator. Battery temperature, especially in nickel-cadmium batteries, should be monitored to prevent overheating, which reduces capacity. A combination of high battery temperature (i.e., in excess of 150•°F) and 7-16
" ,
CEtL VOLTAGE
AT200A, 0 of
AT 2hr IuIRATE
14DAYS AT 80"F
CELLS
AT200A, 800F
lb
0.8I
'AE• O F C H A RG •. N o . O F
CAPABILITY
TYPE
S-
h /
- h / b-
UTILIZATION LOAD ANALYSIS
I he utilization load assignments should be based upon the most dcmand;ng conditions likely to be cncountered during operation. For example, it shoud be assumed that the aircraft is oper, ting at night, with landing lights used during takeoff, climb, and landing. Approximate data conoerning the duration of each load should be known or assumed. To obtain conservative results from the load analysis, the intermi:tent load peaks usually are considered to occur concurrently. Despite the short duration, heavy loading, such as during engine starting, can reduce the battery capacity sharply and should be considered. The load analysis is prepared for an arbitrary set of operating conditions. Too svere a set of conditions would overburdcn the power system during Fpst operations. An overly optimistic choce ot conahtions would limit the usefulness of the aircraft. The
.
)
AMCP7M ALTERNATIVE CHARGING METHODS
TABLE 7-3. METHOD
r
IF EECTONBA-TFRY PERFORMANCE OPERATIONAL
MAINTENANCE
PRINCIPLE OF OPERATION
CHARGERCONSIDERATIONS
SYSTEMCONSIDERATIONS
CHARGED AT CONSTANT BATTERY TRUIECONSTANT POTENTIAL
MAX ELECTROLYTE FAST RECHARGE 055 15LE LOSSEACHCYCLE
VOLTAGE. INITIAL CHARGE CURRENTLIMITED ONLYBy SOURCEANO LINE IMPEDANCE
DOESNOT REGUIRE CE ES MATCHED CLORELY
BATTIRY TEMIPERATU4RE REQUIRES SENSOR
BATTERY CHARGED AT TEMPERTATURE.COMPENSATED VOLTAGE, CURRENTLIMITED INITALLY By CONSIDERATIONS BATTERY
CON;TANI POTEN'IAL/ CONSTANT CURRENT
SIMPLEDESIGN
REQUIRES W1IE RANGEOE: POWERHANJILING CAPABILITIES DOESNOT AM~IRE FACII [TATE$ ZOOO MATCHING OF CELLS MATCHED CLOSELY QUICK RECHARGE REOUIRES CHARGERVOLTAGE TO B1tTTERY CYIARACTERISTICS TO AVOID THERMALRUNAWAYOF NICKEL CADMIUM CELLS
ELECTROLYTE LOSSEACHiCYCIL
REQUIRESBATTERY LOGEPTEMPERATU1TE SENSOR REQ~'IE MULTI-LEVEL. CONSTANT CURPENI
~
TR CHARGE REDUCEOAS STT-CAG NRA;T STTPFCAG NRAE
Ek- IARGE TINE THNONAT OENSTIAL T
EECRLYE ELOSECITROCYCLE YETA OSE~
YIDUCED PUWERCAPABILITY AS OPPOSEDTO C NWANT POTENTIALMETHOD)
CELLS MATCHED REQUIRAES CLOSELYFORCAPACITY
DTIFFICULT TI. DETERMINE OliMuM CHARGE TERMINATION CONDITION BATTERY EPFOIIIRYS SENSCR TEMPE[RAYLIRE
CHARGING CONTROLLEDOR PROGIIAMVIEDSO'HAT EATTLYT RETURNSTO FULLY CHARGED ELECTROL YTE PEARCHNAGE 1PPC)
PULFIE CHARGING
I AS' RECHARGI EORSINEATG
PT AC- BATTERYRfCIURNS WHEN BATTERY FREQUENCYOETERMINED CURACY OF STATE-OF-CHARGE TO FULLY CHARGEE NOT!YIJILY SENSING CZVICE. WHICH MUST CONDIitON CHAkGED; BE RESETPERIOOICALI.Y TO NULL OUT ERIRORS
CHARGERPROVIDES REVERISE CURRENTPULSEAFTEREACH CHARGING PULSETO CAUSE DEPOLARIZATION OF BATTERY ASPLATESTO ALuOWBETTER
ELECTROLYTE LOSSEACHCYCLE
SOF-TION OP CHARGING CURRENT
ORAhINUDVC Q01LIT MONITORING
RQUESCL!MTHD LSEYORAAIT
REQUIIRES COMPENSATION FOR CHARGE EFPFICIENCY ANDI POSSIBLYSTANDRYLUJSSES
SHOULD REDUICE CHARGFTIME SINCEHIGHER POWFll HIGYH CHARGERATES REQUIRES C.YPABILITY CAN BEUSEDAS
REQUIRESCELLSMATCHED CLOSELYFORCAPACITY
COMPAREDTO CONSTANT EQ.-
POSSIBLYAEGUInES BATTERY EPPEC FOREMOST REDESIGN,
POSSIBLE EM, PROBLEM
YTlE__ITII
jSHOULD REDUCE C'HARGETIME
PULSED CONSTANT
CHARGE CONSISTSOF CURRENT ,TJL5ES WHOSEPEAKVALUES MAY ELECTROLYTE BE AS HIGH At BOOA. WHILE CONTROLLING THE DUTY CYCLE LOSSEACH CYCLE TO OBTAINITHEDESIREDAVERI-
AGEVALUE OPCURRENT
I
______
I
choice is, necessarily, a compromise. For most aircraft, the following is sugge~sted: 1. Bittery opcrating temperature, 00 F 2. Duration of flight 3. Night operation. See Fig. 7-11 for a sample set of utilization loads roic. n CUR basd b Es tpialmisin VY EN STAiclmsso R TIN 7-3.4 HE V U RN T R IGinsure REQUIREMENTS s\ome aircraft crigine, have starting characteristics
AS COMPARE TO CON STA:NT) EYETNTiAL
HIGH POWER REQUIRES CAPABILITY
IZATI)N
RQIE EL ACE ACE EL RQIE CLOSELYFOR1CAPACITY PSIL PSIL
M M
RBE RBE -
.
that place severe loads upon batteries, particularly in extreme low-temperature crnvironimenu, or where the battery is not used for long periods. Under such cond i~ons, not only is it more difficult to start the engine, but the battery itself is less active electrochemically, causing the internal resistance to increase greatly. Additional energy must be incorporated in order to a reliable system. This can be accomplished by thec use of parallel-series connected batteries, to sa to provide a marked increase in the available voltage 7-17
AMP 706-202 •- -. f ioq
I
I~
CURRENT
TAKIING
Ln c
'IIIPMN
4
62
62
SIARtLH
1
150
0.5
150
RELAY-GAIlIER
1
0.7
-.C
0
INDICATOR LIGHTS
1
0.5
C
0.5
INSTRUMENTS
1
0.5
C
0.5
0.2
8.4
POSITION LIGHTS
0.7
0.7
0.7
0.7
0.7
0.7
0.7
0.7
().7
0.7
cn -Inacj 0,
0.5
0.5
0.5
0. 5
0.5
0.5
0.5
0.5
0.5
0.5
0.5
0.5b.
0.5
0.5
0.5
0.5
0.b
0.5
0.5
0.5
0.5
0.1.
0.5
0.5
5.0
5.0
5.0
5.0
5.0
5.0
5.0
5.0
5.0
5.0
5.0
5.0
5.0)
3.2
0.2
0.2
0.2
0.2
0.2
0.2
0.2
0.2
0.2
0.2
0.2.
0.2
I(.4
8.4
8.4
8.4
8.4
8.4
8.4
4.51 4.5
4.3
LANDING LIGHTS
1
8.4
MISC. ELECTRICAL CHECK OUTI RADIO RECEIVER
1 Al.0 4.5
C
4.5
4.5
RADIO TRANSMITTERI
1 10.5b
2.0
10.b
10.5
20.0
20.0
Figure 7-1M. Sa.mENpCe -CSIGNIFILtS CONTINUOUS
,n
0.5
C
i
I ''=:
0.5
C
TOTAL UTILIZATION
uCE5
rý
0.7
1.7
1
c:!
o
0.7
0.2
(.l,
unI~C. oE
, .71 0.
3
5.0
cE
-
10
INSTRUMENT LIGHTS
LAIN
TAKLOQEF AND CLIMB In E o~ 'n I InnE1,n5
z
ACtJL
4.5
4.5
4.5
20.0
4.5
4.5
11).5
10.5
4.5
10.5 10.5
.
7.of Ut
VlOzat LoaT AL
o
and current, resulting in a higher power ongreater torque at thease c or necrod initially in parallel and then switched to series afterstart. a predetermined time delay for completion of the
7/4
7-3.5
7-4.1
Battery maintenance can pose a serious o~pc-
4,5
10.5 lo.b
Igme.I.~m~
MAINTENANCE
20.0
,
ormv
Dc should hos provide for cas) acci. to remove o a startell batteri undwt eriesational conditions. Battery installation is de-scribed in par, 7-7.8.
VOLTAGE REGULATION AND REVERSE CURRENT RELAY,: DC VOLTAGE REGULATION
DC aencratina .sy.titmn in raG!! h .elireonle con.•i.t
rational problem, Becausn of the necessity for periodic addition of water to the cell electrolyte, the battery generally is removed from the aircraft on a
of torwith a serias-field starter, volragea starter/gen regulator, reverse-current relay, overvoltage reluy, field relay, starter relay, and start-control relay.""
scheduled basis, and operational delays thus are cncountered, Operating water loss results from the two natural functions of evaporation and electrolytic dissociation. Except at extremely high temperatures, water loss by evaporation can be considered unimportant. Dissociation occurs at a relatively constant rate, and is a function of voltage, current, and temperature. Higher voltage will increase the overcharge current and gaising rate (Fig. 7-12). Higher temperatures will increase losses by evaporation and will lower the potential at which electrolysis occurs. Higher temperatures also will increase the overcharge current when charging at a constant voltage, The requirements for battery maintenance must be considered in locating the battery compartment. The 7-18
One or more of these components may be supplied to the contractor as Government-furnished equipment (GFE), and thus will establish some arc.% of the design. Military Specifications for a generating system using carbon pile regulators and reverse current relays- include MIL-C-5026, MIL-R-6106, MIL-G6162, MIL-R-6809, MIL-R-9221, MIL-R-25078, and MIL-R-26126. 74.1.1 Voltage Regulaior A voltage regulator designed to incorporate all but the line contactor functions, and having additional functions such as feeder fault protection and field weakeiing for shunt starters, is available. The Military Specification for voltage regulators of the
' '"
AMCP 7065202 Wu
j•
-
-
1
-.
,,
"',
S".
• _-
,
' 0 O
~
-
r.'
4 G
vide, as an integral pArt of the startcr/generator. a tachometer generator that enables the iegu!ator to sense speed, thereby terminating the start at a preis weight •dctcrmined speed. economy. The primary advantage of sht-nt I A line contactor. designed in accordance with I
' I "starting I I
So•
•, : .- 4• --Sf
~
.1 VA
0L
CUNkNI
0F D
......
Io i... ,•
t2
14
1b
IRa
__-i-.-j'o
...... , A
Figure 7.12. Gaem Emitted from Nickel-Cadmium Slitered Plate Cell During Owercarge at 70 0-73F
)
-
*
--
static. type is MIL-R-23761, However, as thi% spcification deals only with voltage regulation and paralleling, the system designer must consider the ftiritionlng of voltage regulation. generator paralicling, field weakening for shunt star'ing. line contactor control, engine stirt control, and protection against reverse current, overvoltagc, overc•citation. startup into shorted bus, and fceder fault. It is recommended that, whenevcr possible, the static type voltage regulator be used. Uhe regulator procurement specification should include all of the foregoing funi.tions. This will economize on weight and installation time since separate componcrts will not be required. The switching action of som- static voltage regulWtrs has " app!rc.ti.n probl'0:s. Sw-chinfrequencies that are kept constant, and at values above 1000 Hz, generally will be above any engine or generator resonant frequencies. This switching action also can produce some radio frequency noise; but if proper switching speeds ard filtering are used, radio noise can be held to a minimum. Locating the reguiators close to the generator also will serve in keeping down radiated and conducted interferences. The use of shunt starters with field weakening is a recent approach to turbine engine starting. The regulators sense the voltage on the starter/generator at the equalizer terminal, and use this variablecurrent voltage by varying the shur, field current in the starter/generator so as to provide a predetermined armature current. Starter/generators with interpole 'indings can develop a shunt field current that can result in no-loaa, overspeed self-destruction, In case of shaft failure, a means must be provided to limit the no-load speed. Some manufacturtrs pro-
MIL-R-6l06 and with the proper rating, can be used to connect the ge .crator to the bus. This contactor also would be used fLr the starter armature current during starting, and by tht; reverse current ovcrvoltage, ovtrexcitation, startup into short, and feeder fault functions in disconnecting the generator from the bus.
A relay in the regulator should be used to de-excite the generator in the event of overvoltage, overexcitation. startup into short, and feeder fault conditions. 7-4.1.2 Reverse Current Pelays The reverse current (cutout) relay is designed to connect and disconnect a generator automatically from the bus in a 28 V DC system. The reverse current relay will close when the generator is producinR 18 to 28 V and is at leest 0.5 V above the bus potential. Depending upon the unit-rating, when the generator voltage drops below bus voltage, the relay will open %ith a given reverse current. These units are available in 100-, 300-. and 600-A continuous ratings. Dependin 3 upon gen•erator -apacity, teversc current relays shall be sized to match the maximum continuous generator output. 7-4.1.3 Overvoltoe Relays Overvoltage relays are used to remove the generator from the bus by tripping the field relay if the gornerotor voltaae exceeds a specified limit. 7-4.2 AC VOLTAGE REGULATION MIL-G-21480 is a teprcscntative Military Specification for AC systems. Highly rclipble control units - which provide voltage regulation, field relay control, contactor control, and ovcrvoltagc, under. voltage, feeder fault, and underfrequency protection are available in solid-state versions. AC generator manufacturers design and build static AC voltage regulators to match their generators. I[he designer must consider the electromagnetic interference requirement, regulator operation environmental conditions, and the qualifitation data before choosing a rtgulator. 7-5
OVERLOAD PROTECTION
7-5.1 GENERAL The primary objectives of overload protection are 7-19
4L
to limit malfun~ction automaidcially toa1 single circuit, and to minimize the danger of &moke and fire not only in the components, but also in the wiring. Overload protection of the equipment should be considered separately 'rem circuit overload protaction. In order to obtain maximum safe use of the equipment, any protection required shall be integral. If the equipment is not requked in order to maintan controlled flight, and maximum equipment use is not necessary, the equipment and circuit protection may be accomplished by the same dzvike, provided that this dual function does not conflict with the basic requirement of protecting the wiring bringing powcr to the equipment. The primary intent of circuit protection is to protect the interconnect wiring and the eq~uipment. All load measuring be oroied itha soft form ofe bruitoproe tltoa. sbepropied wileti onefrtheui of protetv-dvc tecton.Proer elecionof he rotetiv deice should result in the lowest rating that will not openl the circuit inadvertently. A circuit-protection device shouldj be umed at any point in the circuit where the wire size cI cges, un-
that responds to a 1mr~aietic effect rather thar. to the heating effect of the current carried by the bicaker. Magnetic circuit breaktrs normally incorporate time delay so as to avoid nuisanca tripping from current surges ef short durat'on. Although the magnetic circuit breakers ame less affected by adverse environmont, they are not used to the extent that thermal circuit breakers are hecause the trip characteristics of magnetic circuit breatkcra may be affected by their mounting position and vibration.
7-.. ktmm Coto Urcailt Brerkers A rtmote control circuit breaker consists of a contactor whose solenoid circuit is coaitrolled by a current-scrnsitive element, plus a manual-switching adtrpiicin We.Teltrui oenosists of a mntia.ally operated circuit breaker arranged so as to trip wlienev,;r the rrmote sensor trips. The remote ciicuit brcaker can be utilized best for bus feedrs and wiring connected to a single load. Although an approved remote control circuit breaker isntailbMltrySefcton ILC833 I--38 pcrcto i o viaiMltr ý# is being developed for a family of remote control cirth ,nara~m ,n,,L~l f Ii. 'h.,y.,i~~tin wire. Where moethan on .rcuit isfed from a sig cuib~~ circuit-protection device, the protection should be 7-51.3 Currtt Ses sizzd to provide adequate protection for the inA current sensor is used in conjuinction with a condividual circuit. The circuit prolection should be lotactor and a manual-switching or trip-indicating decated as close to the power tsource as is practicabic in vice in order to obtain the actuation of a remute conard .r to minimize unprot-ucted wiring, trol circuit breaker. Tht sensor c'lrrent-sensitive cdo7-5.2 OVERL9AF! PROTECTION DEvICES ment controls the solenoid of the contactor. The tripindicating dt.vii.e often consists of a manually Overload protection devices fall into three. cateope'.-ated circuit breaker arranged so as to trip whengoriess circuit breakers, including remote Orcuit ever the current limit of the sensor is exceeded. When breakers; current scnowrs; and fuses. abeaer crcut t ~n~t acurrint 0nmso W 7-5.21 reilr, ~ *r'p-indicating device, the lowest possible rating should be used in order to obtain an immediate iuidictutedeithr termlly Circit ~eaer2n~aybe Circit myeakes beactatedeiter termlly cation of when the sensor has tripped. The current or magnetiwlly. Both typr4 are covertJ by MIL-C sensor can be, utilized best when there is a need to 5809. control a high-currcrit loai, such as in; motor with a low-current ccntrol circuit, and to keep t1,e high7-51.1A1 ThermaW Circuit Breakers current loads to a minimum length. The actuation of thermal ircuit breakers is depcetdent upon a temperature increase in the sensing 752. Fse 75.4Fu cirmtnt which is produced principelly from the load A fuse relies upon the melting of the cureent. current heatiopg. The thermal element will be affected carrying element in order to open the circuit when an by externtl heating or cooling, and must be derated overload occurs. The four basic fuise types art: noror uprated fihom calibration temperature to allow for mal time delay, very fast-acting, and currentfluctuations in am' ient temjlrature. Tb:.majority of limitiag. the circuit breakers used at tiie present li-'ie are of the tingniai type. Each type of fuse is available in a variety of characteristics so as to meet various circuit require7-5.21.2 MixukascCrcult Broakers ments. For a complete listing of characteristics, mdor Magneti circuit broakers use a trip mechanism to MIL-F-23419 and MIL-F-5372. 7-20
______________AMCP
.7-5Z, OVERLOAD PROTECI'ION APPLICATION
possible methods of compatibility correction or alleviation are disacussed.
Circ,.-it breakers arc preferred to fuses. A fuse must be replaced once its current limit has betrn exceeded, and replazcement with an improper size or type is possible. Circuit breakers shou~d hi. grouped in order 'of function or usage, and should be labeled by function for rapid ieifctn.They should be located in a protective panel, or covereii so as to eliminate the possibility of hazard to personnel or contamination by forcign objects. The placement of circuit breakers in the crew area should be avoided. Or.:y those necessary in order to maintain safe flight should be accessible to the flight crew, as any malfunction must be corrected prior to reinstating the circuit. The installation -.~quirements foi fuses and circuit breakers are detailed in MIL- E-7080.
7-6.2 ACCEPTABILITY REQUIREMENTS Unacceptable equipment responses to EMI levels are exhibited as aural, video, or equipmenit malfunctions. In sonme cases, negative aural response can be acceptable if testing indicates that it does not affeet overall mission capability or ilight safety. EMC tests are required to demonstrate control of the electronic intesrference environment. The detailed requirements for these tests shall be specified in the contracto. s control and test plan. See par. 9-li1 AMCP 706.203, for a discussion of the helicopter system NIMC demionstrationi requirements. ý'n t~sting certoin equipment - for example, ordriaoice - f!.ýr u7 efss`rabic response, it is neccasary to itwwuý that 'the systemr functions within a wide safety mo*'gin. Mib~t.ýfry requirements state that an interferev.-e sigrei impressed upon the most critical point of a subsy'ptem must be at least 6 dB (20 -dB for explosives) below !he level that would cause an undcsirif~c resnonee. Items of equipment that directly flight safety, or thut cause or lead to a uisa:-abort or to failure tc~accomplish a mission, arc determining factcis for the safety margin tests iss indicatcd in MIL-E-6051. 7-.INEFR CES CFCAOS 7-. INEFRNEPCFCTIN Military Specifications require that stfficient tests be made of equipment or weapon systems to insure that they are compatible Spocifications and standards applicable to the design reqaiiremenr~s and test procedures necessary to controll th.- electroniý ;nt~rfPr~nrCC onvironmpnt of sk helicopter are Mll.-B-5087, MILLE-6031, NhilL-I 16165, MIL-STD-454, MIL-STD-461, and hilLSTD-462. Iii gentral, the most current spocification in force will be the controlling factor for EMC quulification.
7-6
ELECT'ROMAGNETIC INTERFERENCE (EMI/YMC:)
GEENALaffc-;t 7-61 Electromagnetic compati')iity (EMC) describes the abliity of aircraft electronic/electrical equipment to perform in its intended c;-erational environracnits without suffering or causing unacceptable d'fgr~kdation as a result of unintentional electromiagnaetic radiation or response, i.e., electromagnetic interferenice (EMI1). EMI is generated by a varying electrical or magnetic field. As a result, almost any device carrying electrical current is a possible source of interference, Likewise, within a weapon sysiemn, cach bubsysterii is a potential victimn of a generated interference. In thec course of EMIC qualification of a weapon system, electrical equipment victim response; to interference sources is defined and evaluated. The solution is to control the EMI by reducing the magnitude of interference, isolating the source, or designing the receptor to be 'ess susceptible to the EMI. To achieve a compatible weapon system, the entire environment, fromn intcrcircwit and intersystem to intrasystem. must be considered by following interference specifications and state-of-the-art engineering designs. The samte results can be achieved by several mear.s; and the best solution depends upon the judgment of the cognizant angir'eer, and upon the budget and time allowance of the particular ap-
?
/
706-202
"~plication.
JThis paragraph outlines the design procedures for
the determinatioai of acceptable EMI levels. In, addi. tion, the identification of sources of interference and
746.4 INTER4FERENCE SOURCES Electromagnetic interference originates from either natural or rnantrade~ sources. Natural sources include atmrospheric, precipitation, corona, and lightning dischargec noise. Natural EMI varies randomly with time, geographical area of operations, and seasonal conditions. Thi- type of interference generally affects a broad frequency range in the low-frequiency band. Manmade sources of EMI are either broadband or narrowband generators, and they must be evaluated and bandied separately. Broadband interference diatributes energy over a wide frequency spectrum, and can be either random 7-21
t
AMCP 706-202 or constant in time and amplitude. Typical broadband generators of EMI are motors, switches, power distribution lines, ground currents, pulse circuits, transistors, and capacitors, Narrowband interference is produced by an oscillatory circuit that contains energy only at the frequency of oscillation or its multiples. The output barmonics of a communication transmitter or its internal oscillators arc typical of narrowband EMi. Spurious outputs of a transmitter or receiver can cover a wide range of frequencies and exhibit the characteristics of broadband noise; however, the energy distribution is
defined sharply.
4. Inierference time coincidence, i.e., signal presentation during timea of receptor susceptibility. The complexity of the subsystem, and the number and magnitude of the internal interfcrence sources, determine the choice of protoctive design approachcs. Basic appwoaches to interference reduction within tne helicopter or subsystem include: i. Dftiagn of inherently interference-free cornponents 2. Equipment isolation 3. Cable routing 4. Source suppression
5. Signal point containment and suppression.
Inherent interferences unique to the helicopter can
arise from sources such as the rotating members of the engine, drive shaft, and main and tail rotors.
t
In small- and medium-sized helicopters, radio/ radar operation frequently is hampered seriously by a phenomenon called rotor modulation which creates problems especially in VOR/ILS, ADF, and some communication systems. Rotor modulation interferences arise due to the chopping or reflection of the RF signal by the main rotor. The rotor speed and the nurmber of rotor blades combine to pass a givcn point
7-.5.1 interferencC-free Compoeats All electrical systems shall meet the limits imposed
by the applicable equipment spec-ification, such as MIL-STD-461. These speci!i'zations primarily are concerned with radiation, and with susceptibility to radiation- or conduction-propagated broadband and narrowband interference. Compliance with these specifications represents maximum state-of-the-art interference control. However, the specifications are oroad and do not necemsariiV soive ihe inic.i-eren:ce
resulting in the modulation of the arriving RF signal These distortion perturbat,,ns (amplitudes, cancellations, or harmonics) can set up interference patterns that create navigation system noise, error, and needle oscillation. The interference caa become critical when integrated flight control systems are used, resulting in helicopter oscillation. The expanding use of helicopters in a variety of ei-
problems arising in all systems. If individual borderline component interference sources are not eliminated, compliance with specification limits does not insure that EMC problems will not develop wben the total system degrades from specification limits. 4-6.5.2 Equipment Isolation and Cable Routing Many EMC problems arr. oolved by positioniing eetrncqup ntoruigcbesshthth) pick up or radiate minimal interfereacc. Lccation
considered previously. These interference effects can
and orientation are two importan'.t parameters in pre-
downgrade seriously, or even prievent, a particular
mission capability. Some interference probles arise
__ ._ .. i:ola.i. attnuates with distance. antenna location and oricn-
from atmospheric field charging potentials, precipi(electation charging, corona discharge phenomenon trons accelerated by a strong electrical field around a sharp point), or triboelectric charging potentials (frictional charging as a result of dissimilar material contact)Of these sources, probably the most noticeable effect for EMC qualification will be produced by the tniboelectric charging of helicopter rotating members (engine, transmission, drive shaft, and rotors).
tetion can prevent or reduce EMI. Simple shielding of cables is not always effective, due to the magnitude of interfering signals. In such instances, isolation of equipment cables is necessary. Scparation of high-level from low-leve, cables may be required, depending upon design and space allowances. Signal wires and primary power cabIls may rnquirm separate routing even when terminating at a single connector. If interference is a result of equipmea.t location or cable routing, the following areas should be investi-
7-6.5 INTERFERENCE SUPPRESSION EMI within a subsystem may be divided into four
gated: I. Power and control wiring run separately from
categories: 1. Device signal interference emi3sions 2. Device susceptibility to such signels 3. Transmission path of interfering signals (solid or wave)
signal-carrying wires 2. Audio frequency wir-s run separately from wirec of higher frequency 3. Provisions madt for the right-angle crossing of sensitive circuit cables
7-22
.
S'qk 4 Pro.: wit types used 5. itaxim.urn -patiai separation of antennas or intet fetrencproduc'iog cables 6. Cr]tOaling of nonintcrferinig equipment away
)
ferrous materials will provide shielding above audio
frequencies (electrical fields).
Shielding used to contain interference is dependent primarily upon the attenuation (absorption)
from ý;nowvn interference sources.
of the shield. Reflection loss becomes an properties important consideration for exclusion of interfering
Satirce Suppresslon and SRsceptibity Reduction After using physical isolation aid cable routing to the maximum extent, additional techniques for EMI source and susceptibility reduction include: L. Grounding and bonding 2. Cable and cquipment shielding 3. Filtering, Source suppression is the application of appropriate bypassing, decoupling, or filtering at the source of interference or at a point of maximum susceptibility.
signals. Discontinuities in a shielded enclosure can: provide an entry/exit path for EMI radiation. Ventilation openings, panel meters, access c*vers, dial shafts, or switches are possible EMI containment problem areas. Interference coupling of electronic subsystems can be reduced by careful selection of interconnecting cables. Types of interconnecting cables available to the designer include unshielded wire, twisted pair, shielded wire (single or double), twisted shielded pair, aid coaxial (single or multiple shield). The selection of interconnecting cables to reduce interference coupling and audio crosstalk will be, dependent upon physical isolation of the operating frequency range, and the power and susceptibility level&. In general. a shielded wire provides protection against eietricatl fiellds, whilz the twisted pair reduc-s susceptibility to magnetic fields. To achieve maximum EMI shielding from cnclosures and shielded cables, it is necessary to terminatc them cffectively to the helicopter unipotential ground plane. Both multipoint and single-point ground systems provide certain design features. Single-point grounding (floating shield) may provide the best approach where the possibility of interference coupling with sensitive low-frequency circuits is a matter of concern. When a shielded cable, in
74.5.3
.
MqP 706-202
7-6.5.3.1 Groundlig and Bonding A fundamental requirement for helicopters is the establishment of a well-bonded, low-impedance t all ,,r-...ll! A %n;t, ground pOmni exte•n•d• potential ground plane prevents EMC problems resuiting from unequal ground potentials and ground loop currents, and reduces the possibility of equipment transmitting or rmceiving undesired energy while insuring that shield and filter applications are effective, Bonding refers to the method in which various subsystems or structures are conneced or integrated electrically and mechanically. Bonding avoids the development of electrical potentials between adjacent metallic parts, and provides hl,mogenous flow of radio frequency currents between subsysiteis and structures. MIL-B-5087 provides detail requirements for all bonding aspects of airbornt systems. 7.6.5.3.2 Stdelding A major area of practic.l EMI suppression involves the application of component or cable shielding. Effective use of shielding requires investigation of the interference signals, and of the nature of metallic sl.ielding. The question of whether the source or rt ceptor is prvented from radiating or receiving undesited signals deserves equel attention. Metallic shielding is dependent upon the ;nterfering sgnal component, e.g., the electrical or magnctic field. The lowest frequency for which a desired "shieldingis required normally determines the type of shielding material. High-permeability materials can be used to improve shielding effcctiveness for low-frequency, lowimpedance magnttic fields. Aluminum, copper, or
a sensitive orircit, is ground-d at both ends for the
return circuits, power frequencies in the ground plane can induce audio frequency interference in the signal wires. When electronic and electrical equipment is distributed over large areas, experience has shown that multipoint grounding is superior for RF frequenries. Multipoint grounding involves shield grounding at both ends of all cables, and at all immediate points where the cable runs through equipment. A ,plikation of proper shielding techniques for interference alleviation should be performed in the following areas: I. The radiation source or sensitive component should be installed in a properly bonded metallic housing with limited openings. 2. The magnetic field should be directed away from sensitive components or wiring by use of lowreluctance, high-permeability matarial. 3. Twisted, shielded, or shielded and twisted cable 7-23
should be used for AC and DC power ci,-:uits in order to prevent coupling of super-imposed EMI noise and transients, 4. Two conductor-twisted and -shielded cables should be used for DC signal, control, and audio circuits. Single-point grounding is required. 5. Single- or mustiplc-shield coaxial uable shoull be used for RF circuits. Multipoint grounding is required. 6. Continuity of shielded enclosures is necessary. 7. Shields should be routed through connectors. 8. Minimum-length grourtd returns should be used, and shield insulation from structural members should be insured.
stallation) in the airframe. This includes electronic ccmponents, clectrical relays, electrical power generators, wires, coaxial cabh:s, junction boxes, test connectors, etc., but does not include aircrew control panels and instrument panels. Electrical system installation should be in accordance with MIL-E25499, MIL-E-7080, and as described subsequently.
..
7-.5.3.3 Filtess Filters are used at the outputs of EMI generating sources in order to prevent EMI signal (broadband or narrowband) interference coupling paths. Types of filters utilized for EMI containment ani attenuation include low-pass, high-pass, and band-pass filters, as well as bypass and feedthrough capacitors. Basic filter pao ameters include capacitance, induciMC.x, and mb-aiax. F•cit paraainte, u,.pI.mhwas filtering action by a differant method; i.e., capacitance by short-circuiting, -inductance by opencircuiting, and resistance by dissipation. i-ltcrs should suppress only the :interfering signals.; However, the filter may have an effect upon desired currents necessary to the operation of the equipment. Therefore, an understanding of insertion loss is important to filter applications. In the application of bypass capacitors, the lead length from the capacitor to ground becomes an important factor. Self resonance nullifies the effectiveness of the filter for signals at 1rejueneieS equaito, or .. . greater than, the resonant frequency. Filter containment of EMI can be effectiv; only if the source can he. shielded and isolated from olher internul circuitry, thus preventing the interference from being coupled into other wiring or circuitry within a subsystem. Such coupling may conduct spurious energy to external wiring, or radiate directly from other parts of the unit. Proper bonding must be used in order to prevent interference currents in the ground circuit from shunting the filter element.
7-7 7-7.1
ELECTRICAL SYSTEM INSTALLATION GENERAL
Electrical system instaliatior refers to the installatior. of electrical and electronic oquipmelt (equipment installation) and wire bandle (electri~cal in7-24
7-7.2 EQUIPMENT INSTALLATION During the design of equipment installations, maitainability, reliability, and producibility must be considered from design concept to the production hardware phase. Close attention should be given to the servicing problems tha: might arise with each particular installation, It is not likcly that all electronic components can be made immediately accessible. The service reliability of each must be considered during design of the installation. Factors such as electronic alignment after installation and acccsibility to test points must be considered. If equipment is installed in rows, front row con-ip'nents must be capable of being removed quicHy to provide accessibility tc rear mounted components. Equipment-mounting hardware should consist of not less 'han Number 10 screws, except where vibration isolators are used, in which case the box mounting screws should be no smaller than Namber 10, with the isolator multiple mounting screws no smaller than Number 8. Care must be taken to insure that mounting screws are not hidden behind flanges and protruding portions of neighboring boxes. For easy accessibility, the straight-in approach should be provided for all mounting hardware. Equipment installations involving the placement of electrical receptacles facing bulkheads or other obstructions must allow sufficient room for installation of the wire bundle with a bend radius in accordance .with MAL-W-5088. as well as room tc engage and disengage electrical connectors without darmaging the wires. If possible, electrical terminals on boxes should permit the use of a ratchet-drive socket wrench for wiring installation and removal. Junction boxes must be designed so as to facilitate maintenance and troubleshooting. Access to internal compoaicrts must be such as to permit easy replacement. Th,. locations of internal components must be ider.tified by permanently attached decals. Foreignobjct protective covers must be provided on all junction boxes. On all nonsealed boxes, drain h,'les must be incorporated at the lowest point. Wiring musi be installed neatly, and numbered or color-coded for ease of maintenance, Relays, resistors, small transformers, etc., must be
.
/
%A
-
45 dog
'(A)
(B) Figure 7-13.
-
Permihsbl CIamP Ddumsmiiea
grouped functionally in panels similar to relay panels, tubing, but shoule .ot be covered by the braid or cxtruded outer jacket of the bundle. All components must be located and i4entified by The primary wirv bundle clamps should be of an means of a decal permanently attached to the panel. environmentally compatible type. Nylon clamps are Power contactors must be installed so that the conpermissible in low-tempsature, low-vibration, easily tactor case or box is isolated from the airframe strucaccessible arm. Plastic clamps are not to br use for ture. wire bundle support in arse where a damp failure "Alljunction boxes and panels should have power 91 ,tsam ^ ,har 6 w X;i. k, Air eho.f ci .... ., c,-iuOf and c.tassis grounds emanaung from ui or to interfere with controls. The preferred orienthe electrical connectors, tation of all wire bundle clamps is with the bell (loop) down. The bell should not be turned upward if the 7-73 ELECTRICAL WIRE BUNDLES wire bundle weight threat 'ns to deform the damp. Clamps of the MS 21919 type may be deformed in Basically, there are two types of wire harnesses ordwr to meet special Installation problems by flattenallowable: ing the dlamp bell, as in Fig. 7-13, to a height no 1kw 1. Open-wire bundles, where individual wires are than 3/4 of the original bell height. The mounting tied in bundles and routed through the airframe ears may be bent, but not more than 45 d4, as shown 2. High-density bundles, where an abrasionin Fig. 7-13. resistarm szovering ic braided, extruded, etc., over the entire bundle.
In either case the wire best suited for the particular application must be used; and, when open-wire bundles are used, the wires shall have markings in accordance with MIL-W-5088 and the bundles shall be tied at 3- to 8-in. intervals. Lacing shall be comsatible with the operating environment of the helicopter, Where high density bunweas are used, the bundles must be taped at 8-in. intervals with a thin layer of Teflon tape. An outer abrasion-resistant covering must be braided or extruded over the wire bundle. Tape is not acceptable as an abrasion-res•atant covering except on repair areas or at the ends of a bundle. Tape must never be used as primary insulation. Repairs to high-density bundles should be made by routing a wire external to the abrasionresistant covering. The external wire must have an abrasion resistant covering. Splices are to be coverd ith an ab.asion-resistant material, such as Teflon
! . zaI'q; -7. ZME V NAE • Terminal strips should bN MS 27212 or MIL-T81714 with MS 18029 covers. Torminal strips shall be installed as shown in Fig. 7-14, with the mounting holes isolated, for example, by filling with MIL-A46146 Type I sealant to prevent short circuits to ground. MS 25227 insulating strips may be used in lieu of potting; however, an additional nut must be installed between the insulating strip and the bottom terminal to that there is no resilient material in cornpression with the terminals. A maximum of four terminals shall be used on one atud. MS '5266 boa bars may be uscd between studs to interconnect terminals. When terminals are exposed to the weather - such as in wheel wells - terminals and studs shall be brushed with phenolic resin varnish. Thf wire bundle *Wl be tied to a terminal at each breakout. Ther sell be at least one wire identi7-25
SELF-LOCKING NUT LOCK WASHER FLAT WASHER -.
A---
_44A
MIL-A-46146 TYPE I SEALANT Figure 7-14.
Terminal Strip Installation
fication number visible on each wire without cuttin3 ties. Lacing (or tying) shall be done with single ties. Continuous lacing shall be permitted only in junction boxes and panels. The end studs used for attaching the MS 18029 terminal covers cannot be used for electrical purposes. If two electrical terminals with mounting hardware arc m placed on the end studs, the self-locking feature will not engage in the terminal cover nuts. See par. 7-8.2.1 for a further discussion of terminal blocki;
7-7-6 DOOR HINGE WIRE BUNDLE ROUTING Electrical components mounted on access doors v,ill require routing the wirr? bundles over the door hinges. The wire bundles shall be routed so that they twist instead of bend, i.e., the bundle shall be routed parallel to the hinge for a distance sufficient to allow he bundle to twist. Consideration should be given to using Teflon-cushioned clamps at the twist points to provide added bundle mobility. Added abrasion resistance at the hinge, in the form of vinyl . , T.n tubing mnay be required.
7-7.5
Wire bundles that are exposed to weather and when doors are opened during flight, or abrasion during ground servicing, shall be protected by extra cove.ring (such as braiding or tubing). Weatherexposed braiding shall extend into the connector back shell clamp, but, because of the wattr-wicking properties of the braid, should not extend into potting or connector waterproofing.
ENGINE COMPARTMENT WIRING
The two major installation hazards encountered in 4ngine compartment environments are heat and viI ration. Special attention should be paid to the highvibration environments of engine enclosures. Wire gage shall be a miinimum of 20 in order to reduce strand fatigue breakage. Wire bundle clamps shall be spaced in close proximity so as to prevent wire vibration between clamps and possibie resultant breakage. Crimp-type contacts shall be used in order to Sclumninate strand vibration breakage due to solder capillary action, Wire bundles in low-temperature areas (200°C or lower) of the engine compartment may be in accordance with par. 7-7.3; in higher-temperature areas and on the engine itself, open wire bundles of wire rated at 260*C shall be used. Particular care shall be taken to route all wire bundles away from sharp edges, and around equipment in the engine area to allow extra room for vibration and for structural expansion and contraction due to ambient temperatures and engine thrust. Wire bundles shall be routed and clamped well out of the way for engine change, and design shall take into consideration the use of any necessary installation/removal ground-handling tools. Fire detector elements ghllbe routed, and securely clamped into position, to eliminate crush possibilities during engine change. 7-26
7-7.7 WIRING TO MO I, WNG COMPONENTS Special attention is required when it is necessary t route wiring bundles to components such as actuotors, missile launchets, or electronic components that move during use or storage. These bundles usually flex a number of times and are critical in their operation. The installation should be designed as follows: I. The wire bundle shel bZ clamped firmly to the moving component so that no movement of the wire takes place at the connector or terminal. 2. The wire bundle shall not be under tension at any point in the movement of the equipment. 3. The wire bundle shall be clamped firmly ti the fixed structure at a position whek e if there is any motion, the wires will twist and not bend. 4. The attach point of the fixed structure must be, whenever possiblen at the center of the arc formed by the moving equipment. 5. If the fiued point car not be at th: center of the
GRýOUJNDING PAD
PRIMARY STRUCTURE LOCK WASHER NOTE: BOND ALL PARTS PER MIL-B-5087 Figure 7-15.
) ...
Typical Connection to Grondilag Pad
74
may be required on the slack wire bundle. Vinyl
fined in only one appropriate specification. Environ-
sleeving is not to be used as a substitute for good engineering. Protect:ye tubing should not ride on sharp edges of structure, 7-7.8
BATIFERY INSTALLATION
Batteries shall be installed so that they are readily accessible from the outside of the helicopter. The aircraft connector shall be of the quick-disconnect MNI 25182 type in accordance with MIL-C-18148, and shall be accessible without moving any equipment or reaching around any obstruction. The battery compartment must be located in such an area that battery gas and fumes will not enter the
cockpit or cabin. The battery compartment shall be
)
COMPONEI TS
moving arc, a loop must be made to take up the slack in the wiring. This loop :nust be of sufficient length to insure that the wire bundle is never under noticeable tension. This loop must be self-supporting and selfforming. The seit-supporting feature can he assisted by a preformed spring steel wire woven in, or attached to. the wire bundle. 6. Attention shall be paid to chafing of the wiring, Added protection, such as vinyl or Teflon tubing,
painted with a material resistant to the electrolyte used in the battery. There shall be no oxygen, hydraulic, or flammable lines in the battery compartment. The batter, cables shall be clamped and protected against chafing during installation and removal of the battery. The battery ground cable shall be attached to primary structure that is heavy enough to carry shortcircuit current without damage. A grounding pad, as cshOwn in Fig. 7-15, may be used to increase electrical current capacity.
7-8.1
WIRE
The choice of wire should take into consideration .. cl'z--. -h . .. f t. i,.
not O,,t ,
also the environment in which the wire must operate. The electrical requirements cap be satisfied by the wire current capability; however, the environmental requirement may be compatitzie with the wire *.'mental compatibility will vary depending upon the type of insulating material used. The designer s&Wll assure that the finished diameter of the wire eected
is compatible with the wire scaling ranges of the connector used and compatible with the connector insertion/extraction tool.
.
7-81.1 Wire Imulatihg Materials 7-8.1.1.1 Polyethylew Polyethylene is a commonly used dielectrical material. It is excellent for high-froquency applications. Howev,:r, because of its physical properties, it has definite limitations s an Insulating
material. Polyethylene pgssesis
low abrasion rm-
sistance; the maximum 3afe operating temperature is only 80*C, and it will burn freely in the preence of an open flame. 7-8.1.1.2 Polynylclilorlide Polyvinylchloride (PVC) has physical properties that surpass those of the basic polyethylene. It poswsess greattx abras-an resistance, higher operating temperature limitations, and increased resistance to flame. However, the molecular imbalance of PVC precludes its use at high frequencies, although it 7-27
"
AMCP 022( is excellent in low-frequency applications where reI"
*
sistance to moisture, Rlame, oil, and many acids and alkainesis ~numerous mporan 74.1.3 ~ ~ ,,., ~scription Fluorin~ated ethylene propylene (FEP) demonstrates, excellent electrica: stability over a temperature range of -65* to +230*C, and is suitable for ultrahigh-frequency applications, 74.1.1.4 Polychlorotrlfluonsethyliea eolychloiotrifluoroethylcne, more commonly known as KEL-F, combines many of the advantages of Teflon with a superior resistance to abrasion, thus enabingit b use oasa tin-alle inulaion without any outer covering or mechanical pro. tection. This material is rated for continuous operationthrughthetemeratre ang of-65 to +15thog0tetmprtuerngCf-.*t + 1500C.draulic 7-8.1.1.5 Polyliexamethylene-adipaisldc Polyhexamethylene. adipamide is a readily exru dabe btte plyiid, kownas y is amiy nme dabof noylo.Bimid, betrkof n assbc~ y poor eamlectical of nlon 11"uscof ts clafvel por elctrcal characteristics, it rarely is used as a primary in' asulation on wire. However, it makes an excellent outer coverings wt icn applied over vinyl insulation. Ex truded nylon jackets are tough cnd resistant to &brasion and oil, and have a tendency to increase the temperature sawbility of the piimary insulation,
The Military Specifications for aimrcrft wire a--e too to cover in detail. However, a brief doof some of the more commonly used typos of wire and of the specifications defining tham is giver. to assist in selecting the specification that satsfie& the general requirements. MIL-W-5086 covers PVC-insulated, singleconductor hookup and interconnecting electrical wires made with tin-coated or silver-coated conductors of copper or copper alloy. PVC insulation may be used alone or in combination with outer insulating or protective materials. It is a good general purpose wire, and is available in voltage ratings from 600 to 3000 V and a temperature rmpg of -55* to + I 109C. The wire construction of this specification contains nylon jackets for increased mechanical toughness and resistance to fuels, solvents, and hyfluids. MIL-C-7078 covers single-conductor and multiconductor shielded wire. The basic wire in this specification is MIL-W-508 and MIL-W-81381. MI1L-iW 1818 covers wifVe designed for internal wiring of meters, panels, and electrical and electronic equipment, and requires that such wins, have mini. mum size and weight consistent with service requiremerits. The temperature rating of wire included in this specification ranges to 260*C, with potential ratings of 250 ^,o 3000 V. This wire is primarily a hookup wire, but it may be uted for wiring elcetronic Teraflorathyeneequipment in protected areas of the aircraft. 7-8..1. 74.11.6 etrahaorcthyeeM IL-C-22759 covers fluorocarbon-insulated. Tetrafluoroethylene (TFE), better known as single-conductor electric wire made with tin-coated,
Teflon, is an excellent electrical balance ar~d. applications. TFE offers exceptional electrical, chemcal proertes an thrma ot vaiablein ny othe wiemaeria. isultio TE isultio is rate fo cotinousopertio at200C. ut rmais feilatcygnctemperatures. 74.1.1.7 Dimediyl-slloxane Polymer Better known as silicone rubber, dimethyl-siloxane polymer is finding widespread application as a wire insulation because of its good high-temperature characteristics and low-temperature flexibility. It will withstand 2000C continuously, and can withstand as much as 300"'C for short intervals. However, iii the presence of flame, silicone rubber will burn to a nonconductive ash, which, if held in place, could function as an emergency insulator. Its abrasion resistance is *improved greatly by the addition of a saturated glass braid. Unlike vinyls, polyethylene, and nylon, silicone rubber is a therinosetting plastic. *
74.1.2 Military Wire Spelfkcadioa
7-28
silver-coated, or nickel-coated conductors of copper wires may be polytetrafiuoro~ethyloen, fluorinated thyenepropylene (FEP), or polyvinylidene fluoride. Thefluorocarbon may be used alone, on in cornbnaton ithother insulation materials. This wire is available in a temperature range of 2000C to 2600C, arnd vlaeratings of 600 to 1000 V. MIL-W-7072 covers low-tension, insulated, singleconductor, aluminum wire for aircraft electrical power distribution systems. Aluminum wire usually is used where an appreciable weight saving can be re~alized. MIL-W-81044 covers a variety of construction suitable for airframe and electronic hook-up wire, ineluding flght, medium, and heavy wall insulation thickness and tin- and silverpl'ted-copper conductors. These wires are rated to 5W0 V over a temperature range Of - 53 to +I150C. The insulation consists of crosslinked polyvinylidene fluoride. Improvvd thermal stability is realized through mole-
'r
AMCP 706-202 cular crosslinking of both materi.Js by the high,nergy electronic beam process. These consiructions provide significant space and weight savings while retSining excellent abrasion resistance, MIL-W-25038 covers single wire for electrical use under short-time emergency conditions involving exposure to flame and temperatures of up to 2000*F. This wire is intended for use in circuits where it is ,necessary to maintain the electrical irtegrity of the insulated conductor for 5 min in a 2000*F flame with the operating potential not exceeding 125 V.
Fittings cover a broad area, and include any fixture attaching to a wire. Two basic fittings are terminal strips and connectors. 7-8.2.1 Termhna Strips strips
requirements. Thus, the selection of a connector for a specific application will involve a compromise. M IL-C-5015 covers circular electrical connectors with solder or removable crimp contacts, and accessorics such as protective covers, storage receptacles, strain relief clamps, and potting molds. These connectors arc for use in electronic, electrical power, and control c-ircuits. They have threaded couplings, and may require safety wiring in order to eliminate inadvertent decoupling in high-vibration areas. MIL-C-26482 covers environmental-resistance, quick-disconnect, miniature electrical connectors with solder or removable crimp contacts and accessories. Thes connectors ha'e bayonet couplings and do not require safety wire. MIL-C-83723 covers an environmental-re3isting family of miniature, circular, electrical connectors. These ;:onnectors may have threaded or bayonet
wires. Terminal of two or more ment for a junction ued aya diconcct inapstris e aso strips also may be used as disconrects in ap-
gd inr c MIL-C-28748 covers rectangular rack and panel and electrical connectors with nonremovable solder
"7-8.2 FITTINGS
Termina! strips are used where there is a require-
plications where :'t is impractical to use a connector, or to simplify assembly and maintenance pro-
couplings.
cnat n removable eoal rm contacts. otcs crimp contacts and
or, ts. ly sb nMIL-C-39012 covers the general requirements for connectors used with fiexibie cofrequency radio or 27212 MS is the etrip terminal The standard The temiablrpesth.S ~ ~ ~ ~ ~ ~712 ~ ~ xastndr MIL-T-81714 which consists of a series of threaded axial RF cable. studs retained in a plastic insulating strip. Each terThe designer shall make every effort to select only connectors that provide common termination minal stud will accommodate a maximum of four termethods; i.e., common contacts, common back hardminals; however, a bus bar may be used between ware, and commor, assembly methods and tools wiles four studs in order to allow for more than using MIL-STD-1353 as a guide. having a common junction. The new NAS standard terminal strip, which con-
sists of series of modules rctainei between mounting rails, offers maity advantages over the old style MS terminal strip. MIL-T-81714 covers environmental feedthrough and noni'mr.dilzrough tef ii,,l srifi ips. Fror
7-9
new designs qualified parts shall be in accordance with MIL-T-81714. This type of unit is similar to an •loctrical connector in concept in that it uses a crimp pin, and in insertion-extraction tool for installing the wires, Each terminal strip requirement must be evaluated individually in order to determine which of the type can be used best.
The proper functioning of electronic systems is taking on increased importance in mission effective. ness and flight saf.ty with the development of electronically controlled, automatic flight and engine controls. Thus, the common occurrence of total electrical system failure ft'om lightning strikes is no longer acceptable and a higher degree of static electricity and lightning prottction must be provided for the helicopter in order to ussure reliable, safe, and effective operation over its operational lifetime. One lightning strike can ft expected to occur on a helicopter approximately e-.cry 2500 flight hr (Ref.l), depending upon aircraf zone of operation, mission. normal flight altitudes, susceptibility, etc. Minor to serious structural damage cav result in cases where protection is not provided. New materials, such as pbolyurethane paints. have many advantages relative to corrosion protection;
"74.2.2 Conmetors
)
The ideal situation, as far as reliability is concerned, is to have continuous conductors throughout the entire circuit. However, this usually is not pussi•': interconnects must be added to facilitate assembly and maintenance. The designer must select the connector that best combines high-performance factors with capabilities for meeting env'ronmental
LIGHTNINC( AND STATIC ELECTRICI'TY 7-9.1 GENERAL
7-29
AMCP 706-202 but their exeleknt dielvctri: characteristics also can introduce serious static electricity problems. The high dielectric strength of the painted surface permits the buildup of 5000 to 50,000 V from friction charging of the surface, which may be followed by puncture of the base metal and accompanied by an energy releast in tens of joules. This can cause precipitat'on static or streamer radio interference, and - if the paint is covering an elcctiical component. such as an engine inlet heating grid - also can result in a short circuit of the element. This often is followed by burnup, as a result of energizing of the initial spark by the power system, with resultant major damage. Possible internal problems with high-quality dielectrics include the charging of fluid lines from the liquid flow and the charging of painted internal fuel tank walls from spray electrification or sloshing. 7-9.2 LIGHTNING F OTECTION FOR ELECTRONIC SUBSYSTEMS
The designer of lightning protection for helicopter electronic subsystems should make maximum use of the metallic frame and skin for shielding purposes. Specific lightning protection, or lightning-resistant designs, should be provided at the major lightning entry points. These include main rotor and tail rotor blades, antennas, nrvigation lights, pitot-static tubes, active electrical discharger probe heads, and any other electrical -omponents exposed on the exterior of the helicoptcr. In addition, because of the generally reduced shielding of helicopter frames and skins (compared with fixed-wing aircraft), greater considerations must be givan to magnetic and electrical field penetrations into the vehicle interior. Where all other factors are roughly equal, it i' advisable to use mechanical primary flight controls as engine and rotor controls and to use the electronic systems prin-arily for trim or management controls. Electronic surge suppressors of various types, such as gas or zener diodes and simple capacitors, may be used on critical circuits for suppressing the residual voltage surgc (which can penetrate despite the external lightning protection design), particularly if the electronic systems require very-low-voltage protection. In summary, the preventive design approaches are: I. Principal lightning protection efforts should be directed toward blocking electromagnetic energy entrance through electroma.gnetic windows such as navigation lights and antennas, S. Use of electronic systems for primary flight controls should be avoided. Use should extend only to trim or management. 3. Surge suppressors should be used where 7-30
required, eith(r because of large surge voltages that cannot be reduced at the entry point or for lowsignal-level circuits that require low-level protective devices. 4. Simple lightning test facilities should be used to permit quick evaluation of component performance. Untested lightning protection designs often have proved to be not only ineffective, but sometimcs more dangerous than the components they were intended to protect. Lightning protection through geometrical configuration control of external components, such as antennas and navigation lights, has proven to be one of the most effective methods of preventing lightning penetration into the aircraft. For example, tests of navigation and cofision light designs have shown that a 1-in.change in a cover screw position can reduce the resultant lightning damage from total destruction of the element, with major energy penetrations into the vehicle interior, to negligible physi-
cal damage resulting in voltage pulse amplitude re4;uctions to a few hundred volts. Thus, geometrical control of all external components for lightning proivt;6iun purpubcz gcncraiiv is itc most economicai approach, in terms of weight and cost. Typical entry points requiring protection design effort are shown in Fig. 7-16. Earlier HF and UHF antennas of the voltage-fed type constituted one of "he princip~t electromagne',windows through which lightning energy could enter the vehicle interior. To offset a possible total electrical system loss, these units often can be replaced with shunt-fed antennas, which are inherentl) grounded designs in which the lightning energy essentially is channeled into the external vehicle skin, with only residual high-,oltage, low-energy pulses entering the electronic systems. HF lightning arresters are available commercially for HF antennas. ttid their effectiveness in preventing bothi structural and radio equipment damage has been demonstrated in their use on commercial jet airliners during millions of flight hours. Other external components, such as pitot-static heads and active discharger probe heads, require typical electronic system protection approaches. The pitot-static heads can be protected effectively by conventional electrical system protective devices such as zener diodes or gas diodes; however, the high-voltage active discharge probe heads require more extensive protection development because of high operating voltage levels. For electrical surge suppressi n, mac:y types of devices are available commercially - including zener diodes, gas tubes, simple capacitors, spark gaps, and
AMM 7W
LIGHTNING DISCHARGE
PITOT TUBE
( NCOLLISION LIGHT
(
,
•-
....
•"
I •,q-•UGHr(~ COLLISION
7.
EM FIELD -
I
I-
'
\
\
\ANTENNA
AFT NAVIGATIO(SI LIH
, • ANTENN4A
ELECTROMAGNETIC FIELD PENETRATION THIOUGH PLASTIC COVERS
1
) Figure 7-16.
Typical Lightlng Electrical Circidt Entry Points
silicon controlled rectifiers (SCR). The prinCipal problem in their application ib the selection of the right device, or combination of devices, for the particular equipment being protected. As an example, for antenna front ends, semiconductor devices have a major sh,'wtcorming, th.;y introduce cross-modulation through their inherently noniinear transfer characterutici. Simple gas tubes present only a light additional capacitive load on the front end, and thus provide more suitable protection for this application. For other types of comnponents - such as ciectronic contro! systems, where nonlinearity may not be as important as is obtaining sufficient lowvoltage protection levels - zener diode protection devices may be more suitable. STATIC ELECTRICITY The requirements for control precipitation charging are much mcre severe for htelicopters than for fixed-wing aircraft because of the cargo-handling requirements. Potentials that would be acceptable on fixed-wing aircraft -- 20,000 to 30,000 V, which is well below the radio noise threshold of the vehicle can represent a serious shock hazard to ground personnel unloading cargo from helicopters, and posbly carl cause ignition of ordnance or fuels. Several
7-9.3
general approaches have been suggetd and carried to various degrees of development, including use of active dischargers in which the clectrical field from the aircraft is measured and an opposite charge is ap plied to the vehicle, the use of passive wick-type discharge devices at the blade tips, and the use of conductors hanging from the helicopter to the ground to discharge the vehicle before the ground crew contacts the load. The active dischargers suffer from several disadvantages, including indicating a charge in external electrical crosaficids when using single-head field meters when no charge actually is present on the helicopter and thereby charging the vehicle with the protection device. This can be prevented by using dual field meters, one above and one below the vehicle. However, space-chargc shielding of the field meter sensing head can occur from a recirculating charge during hover. It generally is acknowledged that the use of active dischariers, in spite of the shortcomnings, is advisable, rarticularly when ground handling is frequent. The passive wick dischargers located on the blade tips have the advantage of simplicity, but suffer from the fact that substantial potential is required an the vehicle before they begpn discharging, i.e., they do not 7-31
,
AMP706-202 bring the vehicle potential down to zero. This still permits sufficient potential to give shocks to the ground-handling crew. The technique of using a conducting cord from the vehicle to the ground, and permitting it to contact the ground before the ground crew handles the load, has the disadvantage of the cord being whipped by helicopter downwash. and will not nocesarily hold the vehicle pottntial down continuou.ly while the ground crew is in con tact with the load. The other major problem with external static clectricity on helicopters is radio interference. The complexity of the problem is caused by: the variety of charge-generating mechanisms, of nois-generating mechanisms, and of coupling modes into the communication systems; the difficulty in separating the effects from internally generated equipment interfcrence; and the differences of effects upon different types of equipment. The basic method of controlling radio interferenceincludes: I. Avoidance of all electrically floating external sections on the aircraft 2. Use of some type of active or passive discharge.- in order to reduce the potentials, on the vehicle under friction electrification conditions 3. Location of antennas in areas where the DC electrical fields are minimized under thunderstorm crossfield conditions 4. Use of radio-interite,-,cc-resi,"ant antennas 5. Coating of all external diclectric surfac-s subject to particle impingement with resistive paint&so as to prevent streamer interference, particularly over plastic sections where the interference coupling is most evere. In addition 'o the external problem, which is complicated by the difficulty of proper identification of
the interference source. internal static electricity
7-32
problems involve the fact that helicopiers often are engineered by designers who posse little knowledge of the hazards posed by electrical interference of fuel systems. As an example, plastic tubing often is considered for fuel jettison tubes. Friction electrification of the plastic surfaces of the= tubes can ignite the fuel vapors, particularly when the fuel tanks and jettison tubes are nearly empty. As a solution to this problem, it has been saggested that all dielectrics with a resistivity )f higher than 10' ohm-cm be carefully considered for aircr-,ft use. Thus, the use of such materials would be permitted, but freedom from static electricity hazards would have to be assured for each specific installation. LIGHTNING AND STATIC ELECTRICITY SPECIFICATIONS There are a number of Military Specifications containing reierences to surges and protections. MILSTD-704 defines the accep:able limits of transients on electrical power systems. MIL-A-9094 specifies the requiremerts for aircraft l~ghtning arresters for HIF antennas, and it probably will be extended to indlude all surge penetration into vehicles. MIL-E-6051 is the eiectroinagnetic compatibiiity specification, and refers to permissible EM pulse limits. MiL-B5087 is the standard military bonding specification and covers test current waveforms, bonding jumpcr sizes, protection of canopies, and lightming-induced surge penetration limits. There are other specifications with reference to lightning, but those listed herein are the principal ones with specific data on waveforms, test arrangements, and requirements. .
REFERENCE 1. Rotary Wilng Aircraft Susceptibility, DN 74A,
AFSC DH 1-4, 10 January 1972.
AMCP 706-202
CHAPTER 8 AVIONIC SUBSYSTEMS DESIGN &-I INTRODUCTION
I
8-1. 1 GENERAL clctrni")is efiW astheapAvioics(avatio Avonis (aiaton eectonicj i defnodas te aplication of elenronic techniques io accomplish such functions as communication, navigation, flight )ntrol. identificatio~n, sensivig, surveillance, and terget deaignation. The avionic subsystems will be defined by the detail specification. This chapter will discuss desin rquiemets o iterliic t'eses~asysems weith th uiemheniots toStractcs:sbytm withthe elioter.system. From an operational viewpoint, the helicopter avionic cowplrement can be subdivided into (1) the basic helicopter configuration, and (2) the specialmiwoneqwpent.8. The batsic helicopter configuration as discussed in thic handbook is limited to the space, weight, ad power requiremnents of the minimurn electronics necessary in order to provide the basic mission capability for a specific ciazz of heiicopter. The helicopter classes include light observatio'n, utility, tactical and heavy transport, and external heavy lift transpor-.. Special-misuion equipment is defincd as the additional e,'ectronics - beyond the basic communcation, navigation, and identification functions - requnired to accomplish specific missions such as IFFK flight, night operation under reduced visibility conditions, target detection and recognition, target dcsignation, and integrated fire control, such as is lound in gunships and tactical aircraft weapon systems.
magnetic compatibility/intcrfercnce (EMC/EM 1) must be considered. in general, the ioflowing dcsilbn sequencing must occur. 1. Determine the avionic requirements. 2.Dtrieheaonchrctiscs t 3. CDnteruineth ablonck diagramot:erinterfces heitefcet agam. the eonctrictal syste n h ar system.h 4.eeletricalbsclyoto ntear ai aototesser eeo 4 craft for mock-up purposes. . Develop a schematic wiring diagram for the 6. Develop an interconnect diagram. 7 eeo at it 7 eco at it Develop a wire list. .Dvlpaneetia la nlss 10. Complete a preliminery EMC/EMI analysis pa o h ytm A-i.!
Jmedium
Avionic procurement, installation, and quaiii-
)
cation, Along with bench, preflight and flight test requirements, are defined by Military Specifications such as MIL-STD-454, MIL-STD-461, MIL-STD462, MIL-STD-704, MIL-B-5087, MIL-W-5088, MIL-E.540, MIL-E-6051. and MIL-1-8700. Tbc first step in avionic systcm design is to determine the proper location for each individual system. Because avionic systems are made up of several subsystems and coimponents, it is mandatory that the total helicopter systcrr. and its environmental capability be known. Every avionic system component has temperature and vibration limE*tu!:ons. Before any placement or location is determined, the inter/intra-system compatibility of the location must bi, determined to insure that heat and vibration will not have a detrimental effect upon the performance of the equipment. in addition, electro-
ELECTROMAGNETIC CNPTBLF RGA RGA CMAIIIT
Interference generated by items of electrical/electronic equipment iirutalled in close proximity, as in a typical helicopter systcem, easily can result in an intolerable interference loe, that could reduce seriously the usefulness of airborne equipment, or might even render it ineperative. As defined in par. 9-11.2, AMCP 706-203, the prime contractor shall establish an overall integratcd EMI compatibility program for the helicopter. EMC is achieved by application of an optimum -
mhin,,,i.nn of miannapriai
-*-*----
*-*--
-
anti iephnicM -
-
r~norr-c-
from the earfiist design stage through the: final product or operational feasibility demonstration stage. Accordingly, an EMC program shall be es;ablished that will; i. Insure the efficiprn integration of engineering. management, and q, l1ity assurance tasks as thty relate to EMC. 2. 1isure the efficient integration of EMC withk all other systems an~d subsystems. The first requirement for acnieving EMC in an avionic system is that all major components and subsystems be designed, constructed, -,ndtested in comnpliance with MIL-STD-461l The second requirement is compliance with MIILE-6051 as an operating helicopter system, with all avionics and other equipme~nt installed and per. for-ning their normal functions. B-1
AMCP "&6202j 9-1.3 DEA2GN CONSIDERATIONS The design considerat'ons that follbw are applicable to EMI and should be ased to assist in keeping EMI to a minimum. The first design consideration iinvolves the creation of a good, basic grou-id plane. This is normally the avioic ompnentchasisor te arfrme srucure for the avionic system installation. An ideal ground zro-ipednce plan awileropotetia, povid reference b~se for all circuits, and a sink or trap for all sigalstha ndeire canoecrneintefernce sources.unifornn, desgnurcecsl.l Asecond deinconridecration, patclryat th lower communication frequencies, is the requirement for single-point 1grounding so as to avoid ground loops. The h-ige. circul~ting utirrents in ground loops are potential causes of interference. A third design consideradion concerns shielding practices for major components ard for the tntal aircraft installation.thdeinpaetocne A fourth design considcration calls for isolating, as fras possible, the power-carrying -vires and %cables from the high-impedance, low-lcvel signal wiring. The baý,ic p-inciple is to categorize conductors on the -l their Primary leaakange f basicc _r -h..eth......poinents are mnagnectic or clctrusiatic. Ail condu~ctors carlyin,3 power or signal eniergy have associated withý ar~ external or- leakage field that ran hiduce ut, wanted signals or noise in nearby conductors by 'i ductive or capacitive coupling. To minimize these undesirat~le field components, various techniques are usedI - such as electrostatic and magnetic shielding, spac searaiontwitin of irepais, cossver spaetistng spartio, f wre air, cossver wiring methods, use of field-absofbing materials, and netralzaton ethds.much sophstiate '~
Vthem
A fifth deinconsideration is to provide adequate troni-. eq.uipment, and for parts of tile vehicle structure that can contribute to the generation of clectri cal ncisi. All electrical and avionic equipment, subsystems, and systems that produce electromagnetic energy shall be installed to provide a continuous lowimpclanv'ý path from the equipment enclosure to the aircraft structure. The. designer mi.st demonstrate that thac proposed bon,' 4 methods result in a D-C resistance as specified Jh. ~.i various cl asses of bonding in MIL-B-5087. The design shall minimize the long-term effects of ope:ra' oal vibrotion, the effects of cerjrosion bz~ween adj..- nt surface and of' galvanic aztion, the diC!eCtric breakdown of insulating finishes, and the undzsii'nblc cifect of intermittent electrical contact. Bonding stiali be accomplished by direct mn..tal-to-metall contact whercvcw practicable. A bonding .imper shall be used where direct meta!-to-
~,8-2
mettal contact is impracticable. Such jumpers shall be pcfe nMLB58,o o h appropriate tnadt~Sa types, and shall be kept as short other
r addrctspoibeWhrpatcbltejm hjme addrc spsil.Weepatcbe shlr o xcd3i.i egh ufc rprto for bondý and grounds shall be accomplished by removing all anodic film, grease, paint and lacquer, or othecr high-resistance: materials from the imnmediate area of contact. Direct-to-basicý structure bonding "hlused wherever possible. For vehicles with metallic skin, the skin &Wal be designot. so that a low-impedance skir. is produced through inherent RF bonding during construction. RF bondint must oc accomplished bctwoer all structural comnportents. Hatches, access doors, and similar comnwirientshnot einhe roxuiy tointrfpermanenl soucn-o wiigsalbethrbnetorprantyisulated from the vehicle skin except for the protciesticdanbd.tishgldsrbedin thedstaign phasi, toanfer reishghlaly weithleauirngm eirlwthifam designers so as to resolve compatibility problems. For i;,uidelines to analysis and design, the design er.H14 gne hudcnutMLB58,AS Re.I. and-4cN6S1.douetsrfrecdi I..SD41 A sixth design consideration for rninimniziiie EN.1 is to separate :alid isolate pulse device and equipment from other devices that are high)) isusceptible to EMI. This is accomplished by attemptiall to separate use aas itroaortas sc tesa aas nergtrtas sc tm spit ponders, ani I H F transmitters from com~puters, data processors, and sliscrptible receivers. This is not always possible, -nasmuch as the physical locations of somie devicus are dictated by m.ission requirements. However, the designer should strive to achieve as physical and electrical isolation as is practicable. A fl-!inaldedrg conadialrabls tien cablvteuse o ie f doquilerinhieldescohaxil cables Ote cablesm or 9ire alhveamimmo90 reuin shld providg.Cnncosued withblc shilsfifselin cables shilllde. prvddwtblcshlsfiatengalehcd. -. ENIOM TA AS CS SET 8-. NIOMNA Ervironmental considerations are pertinent to the design of the b,ýsic avionic system, and to the airfrume-systcm interface. Susceptibility to rotor mnodulation must be consdte6e. The very high frequcncy omni-dircctional range (VOR), instrument landing system (ILS) localizer and glidescope, VHFFM %omer, and other cquipn~ent have been affected adversely by near-frequency rotor modulation. As rotor blades pass over the air-.raft, a modulation of the incoming wavefrorit is set up, with pronounced
K
<
~ ZuJ.~ -
.
'S~.j~i.-i *J~k
4;
AMACP 706-202
)
results. In addition, the modulation is in a nonsinusoidal manner and the harmonic content is high. The variable and reference modulation in a VOR is 30 Hz and thc localizer and glidescope frequencies are 90 and 150 Hz; these are all convenient harmonics. This, coupled with the fact that helicopter rotor speed often is such as to give harmonics of 30. 90. and 150 Hz creates prublek-as for avionic system designers. Techniques have been developed, as discussed in par. 8-3.2, for phase inversion and cancellation of the modulation. This technique shows promise of solving the problem. However, the characteristics of certain existing ground "aci~ities are such that when this techniqusv is used one error sitaply is exchanged for another. In any case, it is essential that nmanufacturers, of equipment for helicopters incorporate very narrow bana filters into their equipment, and
based upon heat rise can be developed for various critical ambient tenmperature situations. It must be reanari o vial ntegon or when the helicopter is hovering; therefore, either auxiliary air from outside the aircraft or engine bloed air must be used to meet cooling requirmecnts. if a particular helicopter is to be a multiuse aircraft, requirements applicable to the various uses must be considered. If a single airframe is to be used for two or more different types of missions, the avionic configuration design must accomplish all requirements economically. Environmental test requirements should be foiiiiulated during the initial helicopter planning phase and should consider all enviro.nmental conditions to be encountered.
p'-ocurement should be based upon this criterion. In
8-2
addition, antennas must be decoupled from the main rotor insofar as is possible. In general, helicopter avionics do not have to be desit~ned to withstand the altitude er~tremes that fixed-wing avionic systems do. Dust and sand are more of a problem to helicopter avionic equipmcri whan to fixcd-wing Cquipmontn. Diiit can pack up in voltage regulators, iotating equipment, relays, switches, and othei critical devices and cause malfunctioas, If these components cannot be hermetically scaled, they must have: shrouds or other protective covering, and they will requirt additional mrintenance, Helie'opter vibration must be considered bas'4 upon the installation of all equipment. The addition of equipinent, particularly in the it~struriicnt panel, affects the frequencies and amplitudes of vibrietion. With the &aJvcntof solid-state components, vibration prob~lems have beecn reduced, but most internal Comnponents sitill are, vulrerable to vibration fatigue. While avionic manufacturers qualify their products to a specification, the applicable specification cannot duplicate absoluttly the situations enc ountered in the actual in.,taliation. Therefore, it is desirable to dcsign for a minimum level of vibration. Temperature ranges may be severe. dependine upon ambiznt conditions. Because of the large areas of trinsparent. windihielO or canopy, solar heat can become a problem when the doors are closed and the helicopter is on the groanrd. Avionic packages are good lieo.' sinks; they will absorb a great deal of heat and will kot dissipate it for some .imc after becoming airborne. Some ground tcniperatur! conditions 4re worse than conditions in flight. A temperature survey it an important design considcration for avionic equipment installation. Data
COMMUNICATION EQUIPMENT
8-2.1 GENERAL Army helicopters coutairt many combinations of communication equipment. Because of the rapid de. velopment of new devices and the chanzges in nouierclature, no specific radios are referenced in this chceLer.
The typts of communication eq'ripmen! currently in use include high frequercy, HF (3-39 MHz) -very high frequency, VHF (FM') (30-75.9 MHz); VHF (A.M) (118-150 MHz); ultra high frequency, UHF (AM) (225.400 MHz); and millimetei wave. A typical communication block diagram is shown in Fig. 8-I. Guidelines for the radio installaiwn iaicludcl: 1. Tlae transmitter should be mounted in close proximity to the antenna in eirder to preclude line losses. 2. The control head shoulti bv rr!i -.;r.: I te. provide ease of access for the flight clew. 3. Routing of audio wires should be such asIQpe vent crosstalk and feedback. 4. Power leads should be of sufficient size to permit fUl generator/battery voltage to appear at the radio under transmit conditions. 5. Componenns should be mounted in an area wheic sufficient cooling will be avaik~ble. 6. A low-vibration area should b-- provided for mounting. The vibration limitations for communication equipment are identified normally in the applicable equipment installation specification (SCL-IOOXX). Each of these guidelines contributes to die reliability of the overall system, and, thicieforrr, is cssential lor helicopter mission accomplishment. Transmission lines, usually coaxial cables, are used to cai r) the transmitted signal to the antenna and the 8-3
Iebrdta I
AMCP 706.202 received signal from the antenna to the receiver. Lowloss cable, such as RG-214/U, should be considered for lengthy runs where exmessvc loss could occur. Newer cables are being developed, and appear promising. Commercial cables, even if not yet ap-0 proved by Army qualification tests, should be proposed by the contiavtor if their use assists in maintaining efficiency and low cost. AMCP 706-125 (Ref. 2) should be consulted for further information on transmission lines. Antenna considerations for communication equipment are presented in p~ar. 8.5. 8-2.2 MICROPHONE-HEADSET A helicopter microphone must be of the noise cancolling type; Armny and civiiian experience has shown that a dynamic microphone is the mos~t o.f fective. In the noisecancellinS microphonc, ambicnt acoustic noise enters both sides of the microphone wv~th equal intensity and at the same phase relationship. Unfortunately, the face, lips, teeth, and protective helmet have a major effect upon the noise cancelling characteristics. In tie case of a helicopter with a high ambient acoustic noise level, it may bse nccmary to conduct a power spectral density mmaurement of thc noise love! in the microphone area, us'ng standard microphones and then to develkp a filter that attenuates unwanted noise while permnitting a voice to psi, through the microphoute amp~lifier. The heuadset nlso must be of a dynami~ic typie, ead i: is highly dm~i' able that it have minim'um high-level distortion. Yhe @er muffs should tic lagge enough to exclude c~tr&ncous noise whilc providing optrator comfort on long flikhts. I bt microephonte'heedscet. if included as part of ihe hel~met, also should possess the foregoing characweistics.
O
6 16 F6866666 4 1
MIKE
___
a
1
ICs IESLCTKOFF9
NAy
NAV
VOL.ME
N HOTMIKE
Flgs: 2 W..Typical Iuercomnuulcadoin Solocter Box 8-2.3 INTERCOMMUNICATION SELECTOR Box The microphonc-heidset plugs into an intercommunication selector box (ICS), which will have a number of microphone and headset selector switches. Us5ually, the zekoctor unit has an integral volume controi and a "hot" microphone switch. A typical unit is shown in Fig. 8-2. The ICS control unit swvlocts - for each crew member - the radio for which the hieAdset is seclected, the intercom functions, and the emergency radio. (Th.; cien~criccy radio function noa anjally is. unswitched audio which cannot be disabled.) Often,. other fun-.ti ns are routed through the ICS control box - such as engline failure, warning, landinig Sear warning, rotor brake-n warning, and other audio warning signals. ',he ICS control box is the main bwitchboard for the flight crew. It usually is arranged in a standard configuration so that a craw member can transfer from one type of aircraft to another without confusion. It is imperative that each crew member ka.-ve ready access to his ICS control unit. Cockpit and crew station ar-rangement is described in p~ar. 123-2. L 8-3 NAVIGATIONAL EQUIPMENT 0-3.1 GENE~RAL The. categories of navigational tcquipment are: 1. Terminal maneuvering
Ea route navigation
Becauxe the usefulness of VFR aircraft has been ~limited
Figurt 841. Blcvls WfAgwm of Classical Commssancatkom Sy.e 8-4
6 NAV
Interdiction
HLO~fl3.
L svl4sT...
MIKE
3
2
ANTENNA2.
~
2
in recent crimbat operations, future helicopter icquirements will include all-weather operation with stability augrficntation systems. The location and installation of the anternais required for the navigational equipment dircussed are furthtr defincKA in par. 8-5. Navigation displays arc discussed, along with other flight instruments, in Chapter 10.
\
6-32 TERMINAL MANEUVERING EQUIPMENT
*
*
__
In addition to the basic flight instruments, a helicopter may have VOR, ILS, marker beacon, and, preferably, a radar altimeter. The helicopter VOR Antenna Array is mounttd for horiontal polanzaon in an area that provides maximum performance and minimum rotor modulation. The basic ground signal -is both frequencyand amplitude-modulated at 30 Hz, and the aircraft VOR re~cvr measures the phase difference betwcen theie two) signals in ardor to obtain the angular displacement to the station; accuracy of the VOR is about *2 deg. Because the main rotor blade often ro-tates at an angular rate that is some subharmoi;.of 3Q Hz, helicopter manufacturers have experienccl a great deal oftdifficulty with rotor modulation of the VOR. Usually ftherotor modulation can be seen in the oscillation of the course needle. In extreme prw the course information will become unusable. Ther exerimnta ar som tecniqes aailble by which thi rotor modulation can bc eliminated,
Ewenually. these oc-_n4 upo~n actecting the rotor \amplitud4:-muodulation on the subcarncer, inverting it,
summing'it .with the modulation on the VOR rn effect, canrcceiver variablu -phase channel
,and
I
.3.33 EN ROUTE NAVIGATION EQUIPMENT Navigational equipment used in eo route flying may include ADF, DME, TACAN. LORAN, conaDoprrdandietlnviainyas The missojn requirements dictate the AV='m of sophisticwUon required. 53.1Ato kDieimFae(DF This -is a refinement of tht. old radio direetion finder, Rnd employs both a sense antenna and a ho antenna on-board the aircraft. The uignals from these two antennas are added vectorially and a cardoidr.. pattern results. By means, of circuitry w.ithin the radio, the system always seeks the null of the cardioid and, by means of proper calibration tnd instrumentation, a pointer shows the re'lative bearing to the station to which*the radio is tvned. The equipment 'do~ikner must be certain that dte receive .r antbnnas arc selective, so that extraneous signals will not affect the operation of the unit. Attempts have been mad,; to install ~vidcband amplifiers in the loop and sense antenna circuits in order to tyt mrv h fiinc fasotatnao alemn t usually fain ha antedn
inraddiion how a vaigs
atepsualyhvrsuediadtonldin probletns. Thi'kns 'antenna _sh66ld be as iar from' the main tpevtin-fecerm oW Fsilinrd closerfeoenhe efrobpevn
ceiling it out. While this principle has been demon-
an ihu ise, od tritdeetric no
strted hs benfoud~t italo inroucean din the: spinning tone wheel wheel-tooth syn~metry typiall t ge~rae usd th sucarier rnVOR
trical CG of the helicopter for good reversal &hiat rifc.Tev locato oftesne nen
~ditional error into the system. The small variation in
termiihes ih revrsal chir'cteristics of the system. stemicaresulsedt iencan a elthe moulcatrior ofVc Empirically. it has bee.n gh'own that if the sense antenubsearrier.lt Whn themliud modulation of sbare the no is located forwa~rd on 'the aircraft belly, an early reversal (and. possibly, multiples) will occur, if it is withthevariblephas sinltherot hs sgath oo summed wit th aibl cmponent is cancelled but these asymmetrits imaft on the top,alte ndi fr revesalt will oiurs mounted poea new AM component on the tone. Thus, the thatoptin sfton mounted ocresultdif elyreversal will be applied to operational eqyipmrnpet (Ref. 3). t ',
systeisl toute lanin (nteS) Anuide sthuenarcat toathengrosnd.eTheILS iscopsed to Suid t th th aicrat grond.TheILSis ompsed of a loaiewhich operates betwecn 18ad 12 MHz and uses the same antanna a& the VOR, and a glidescope, which operator. between 329 and 335 MHz. When a localizer frzquency is dialed, the glide-
scept is channeled automatically to the proper fryquency. Both the localizer and the glidtscopc t4e 90 and 150 Hz to provide right-left or up-down sius Again, rotor modulation has been i. major prob'em inboth
\
these devices.
An additional radio for use with the VOR and ILS is the marker beacon receiver. The receiver operates at 75 MHz and is used to locate points along the ILS path.
tset meroa intiti symmercli the longitudinal axis fo'- symmetry in calibration. Because the loop is affected by large masses, consideration should he given to the location of deployable or disposable stores when placing the loop. The- ADF never is to be considered as precision
equipment. It is versatile, and its angular error vanies invctsely with distance and increases with atmospheric noise leval. It is vulnerable to countermeastres. 8-3.3.2 Distance-measring Equilpment (DME) Distance-mcasurii~g equipmnt.r has been in %:sofor some years and is very accurate. It consists of an airborne transponder that sends out a signal that 8-5
IqF
anotd" Signal to the aircraft on a slightly different fnMuqum7c. Th ib. "t~tPnWmasrste tota elapsed time, divides by two, and couve :u this Igua* into miles. The distance thea is presented to the pilot by moma of a dial Wunstinet or a digital display. The antenna should be isolated from other an. tuina as much as possible due to the pulsed Obaractrautic of theoutput.
64.3.3 TastWa Air Navlgatiuim (TACAN)
t iliary avigtio) is emiliary ir Naigaton) (Tctial TACA TACA Ai (Taticl system that combines DME and a form of VOR (Station Waring) 6o as to give the pilot acontinuous P051ion fix with respect to a single station, in terins of distance and bearing to the station. Each ground TAuCAN beacon consists of a transmitter and an antenra. The transmitter operates in the UHF band, between 962 and 1213 MHz. Tite airborne rejcivrtianamitter send& out a chain of interrogation pulmc and deodes the rply from the ground station TACAN is & line-f-sight systeim, end there is almost no chance uf interference from station: beo. Yond tke radio horizon. 6-.3.3.
LAOR-rW116 14aragam11
tMLAIR~N)
Hyperbolic navi1gation is achieved when synchronizett signals having a known velocity cf propagation are tral emitted from at least three known points, and thiL relative times of arrival cf these signals ane measuredl and interpreted, Standard WILRAN is a hyperbolic navigation syste that was developed primarily fer long-range nuvigatsion over water. It operates on one of several frequen~cies between 1700 and 2000 Hz,ý and its propagation characteristics are determined primarily by soil con ductivity and ionospheric conditions. The long pulse length requires the use of rareful matching techniquts in order to achieve re*sonlible precision The chief disadvantages of LORtN indude the impossibility of instantaneous fixing wih out dual insts'altions. the presence at night of long tra~ns of pulses reflected from the ic-nosphere, and the faut that ionospheric iransmission is not homo. genous, so that the shapes of the sky-wave pulses often are distorted and difficult to match, LORAN C/D is ths latest inodei of this type of airborne equipment. The accuracy is extreately good when the set it operating within the range of highest accuracy of the transmitting stations. The LORAN equipment operates in the HF region, so a relatively long, wire antenna is desirable. Except for reversal characteristics, the same considerations should be used in the placement of this antenna as arc used for the ADF sense antenna.
There usually are two types of corupawesyilasm ara helicopter. The most prod"ie s the slaved directional gyro (D%) that has two modes of operation - free directional gyro and slawA~ coampass. In the free mode, it acts only as a dirotina gyro; in the slaved mode. signals from a flux valve i"lve it to mogneuic north. Particular care must be. uted- in the location of the flux valve; it must besas farmsW praoticable front any ferrous material, and any DC wiring in close proximity must be two-wire twiste. The stand-by magnetic compass usually is located above the instrument panel. Compensation is intogral to the unit. If light wires are installed for thit com~pas, they rinuat he two-wire twisted. Intermittent fields should be avoided for all cotnpassses 8--3.36 Doppler Naylgatlea Systems Doppler systems measure velocity only, by the welkonDpercf inhchadtonrma worell-nmowin roplaier efetoihewhich radispation frn ei oinreaiet h iw~ sdslcdi frequency. In practice, this means comparing the frequency of the returned echo with a stable reference frequency; thn6cierence: between the two is a gin=c measure of the relative velocity. Accurecy thus do-( pends upon the echo quality. Echo quality from water, for example, often is poor. Dop~pler systems dettermine location relative to the point of flight origin by integration of measured veolctvcor.Dppeacrcyersesaniprovement over airspeed-clock-compass dead reckoninX because the velocity vectors measured are relative to the ground. Generally, the vectors in the direction of flight and normal to it (x and y) ate Mmasured. The system accuracy is expressed as a percentage of tho distance travelled, as opposed to inertial systems wvhose accuracy is relative to the tine of flight. Typical performance accuracy of Doppler systems for ground speed is 0.11 %(rinD) over 10 nmi, and 0.06 deg (ims) for average drift. Reliability of actual installations has typically been 1000 to 1200 hr MTBF. Further iniormation on Doppler systems can be found in Ref. 4. 8-3.3.7 Inertial Navlgadoa System. Inertial navigation systems (INS) have beow developed primarily for use on fixed-wirg military and commercial virciraft. Because of the requirement that an inertial platformi be precise without ground iation correction, this device is a valuable navigatioaal standard in forward areas. Couniermeasurts arc virtually noisexistent.
9
AMCP 706-2022
INS oporatiork is governed by two basic physicuii priciple - lb. gravitational pull of the earth and the gyroscopic principle. Essentially, the INS is viade up of a stable platform, a computer, a musnory, and a presentation. Many outputs t~an be derived from such a unit - e4g., true. north, velocity and direction, -%rab angle. and all autopilot signals necessary to pmedetermine flogt track. The only requirement necessary for installation of inertial navigation equipment is the provision of a precise iongiudinol Wis for reference. Normal EMliEMC procautions also must be taken.
scaInner,
8-3.4 JNTERDICFION EQUIPMENT Because of security considerations, this discussion necessarily is limited. Generally, a specific electronic countermeasure (ECM) device will live quadrature information (general location of the equipment under surveillance), frequency, pulse width (if pulsed), duty cycle, peak and everage power, and repetition rate. The ECM also is designed to provide other types of information i! required. ECM is desirable as -.nterdiclion equipment, not only to provide the helicopter
8-3.6 STATION-KEEPING EQUIPMEN4T Thie c~uipmcsnt geinerally is used for maintaining a position directli 'wer a point on the earth and for formation flying. The systems usually are employod on larger, load-carring helicopters requiring precision positioning for pickups and drops. In applications requiring loading and unloading from a hover, Doppler radar is the most rcadily aviiable andl precise type of station-keeping equiprnent. Most conventional Doppler navigational
crew with information. but to telemeter the informabion back to secondary forward anaiysis armas. In addition to ECM, it roay be desirable to include optical or laser ranos..flsders, ranging gunlaying radar, or other devices to aid the interdiction aircraft in performing its mission and to pinpoint targets for forward ground artillery. Communications, usually secre, will form a part of the systent. While not actually a part of the interdiction equipsnout, the electrical characuteistics of the helicopter must be considered to be a part of the mission. Acoustic noise, radar reflctivity, and infrared (IR) sianatarp moist hp minimived In sditio;dn, th- aircraft mssonwilspecify radio frequenc~y tranamissian owa.
8-3.5 L0W-LIGHT-LEVEL NAVIGATIONAL EQUIPMENT
)
This equipoent is unique. and is required only for specific missions. There are three basic types of such equipment. The first type is low-light-leve! television. In is simplest foirm it is nothing more than a dlosed-circuit TV employing a c~mea that is sensitive. partic~ularly to low light levels. The second type is an adaptation of the photonultiplier or 'mnouperacope 9' device used during World War 11. Optical stabilization and intensi&ictiontechniques have bein refined, and the imnproved system has some unique advantages, The third, mad most prosnisiag, system is the infhare (13.) detection type. It consists of a sensor--
a signal conditioner, a power supply, and aa video display. The scnsor responds to a selected spectrum in the 1K region, and operates in total darkness. Rt is extremely sensitive and can discriminatea bttween sight temperature differences. The system cao ht designed to include very accurate definition. and thec resultant display on a dark night can duplicato a daylight TV picture. These low-li~ht-lcvel devices are used in activities that caii for radio silence, acoustic silence, low radar reflectivity, znd low iR signature. Flight techniques also are iinportat.
nsvteniq
arre
nsMa an S*dinett tn d-n
rnivip na'ngnti. n-
the frequencies are rlatively high 4nd accurate. At zcro velocity. the Doppler shift will be zero, and determination of movement is difficult; however, recent developments by avionic manufacturers effectively have permitted zero error during hover. This type of system is recommended for station-keeping for loading situations. Doppler syttems require specific antennii locations. Generally, three- or four-beam patterns art used. Some manufaciurers incorporate all functions into one antenna. Certain types of equipment require tha antnn 5 ib~dlruibizunp poses. Provisions for antenna location must be made during the early dtsign phase in order to maximize ef. ficient use of tht area in the fuselage belly. For station-kaeping during formation flying, many techniques have been used in the past and are satisfactory under both VFR and IFR conditions. Cost-effectivenecss decisions will determine equipment selection. Radar, together with beacons, LORAN, or special IR, may be considered.
8-4
FIRE CONTROL EQUIPMENT
6-4.1 GENERAL The airborne fire control system selectively performs the Weks of (1)establishing that the weapon is aligned Properly to hit tie target, and (2) driving and holdinS the weapon platform to a commanded Pon'Lion. 8-71
1h
.1
The major elements of any fire cositrol system consit of sight, sensors, and computei. The weapon controls are a part of the fire control equipment, and their functions are to nctivate the gun or missile, regulate gun firing rate, select the weapon, regulate the ammunition feed system, inventory the ammunition supply, etc. The complexity and sophisticntion of an avionic fire control system wvill vary according to the degree of accuracy required, rnd the type and flexibility of the armament subaysate,. The armament subsystem may be an integral subsy3tem of the helicopter, or it may be a modular comporeeni that can be snapped on or off to suit the particular mission requirements. Thus, the designer must establish fire control design requirements commensurate with required aircraft missions. Because the kinds of missions to b-. derformcd arc likely to be broad in scope (ranging, perhaps, from
weapon is aimed and fired. An optical sight, either direct-viewing or periscopic, Senorally is used for daylight operations, and may provide selectable dogreos of optical magnification. The sighting station also may provide target range and image intensification sensors. Target range equipment can include lasers, radar, or stadiametric ranging devices. Image intensifiers include low-lightlevel television, electronic image amplifiers, and optical telescopes. Direct viewing weapon sights have been a major source of fatal and serious head trauma during crashes in US military aircraft during World War II, Korea, and Vietnam. It is essential that all sighting devices be designed to eliminate their potential as injury producers. Factors to be taken into consideration in order to delethalize sighting devices are: 1. Ability to instantly remove, jettison, or stow during emergency
close tactical support to rescue), the fire control requirements likewise will be varied. Mission analysis will determine the fire control functions to be performed, and a careful selection of multiuse armsmernt subsystem equipment will reduce weight and
2. Not to create additional hazard(s) to other crew pers.nmnel in the event of emergency 3. Adequate stowed tiedown strength to prevent sight from rebounding during impact 4. Not to represent lethal missile 1azard in the
for helicopters include flexible turreted Suns, fixed guns, rockets, and missiles. The interface characteristics of their supporting functions are as different 3S the armament systems themselves. Consequenfly, detailed integration design specifications for each typo of armament subsytem shall be issued so as to insure ffective weapon delivery. -84.2 INSTALLATION Adequate provisions for installation of the elements of the respective fire control systems should be incorporated into the helicopter to insure proper matching or harmonization of such systems with armament. The fire control system should be installed as specified in the helicopter specification governing control of guns, rockets, and guided missiles. The helicopter manuuacturer is responsible for the shock mounting of all fire control equipment installed, Vibration-isolating mounts should be incorporated so that equipment will not be affected adversely by vibrations in the helicopter. Testing shall be in accordance with MIL-STD-810.
*
I
S
8-4.3 SIGHTING STATION The sighting station providts the means by which the weapon operator establishtx the xLý.nuth and depression coordinates of the target relative to the aircraft position. For flexible weapons, the sihting station includes the operating controls by which the 84e•tigb
5. Review of Ref. 5. 8-44 SENSORS Fire control system sensors provide the information necessary for solving the fire control probIen and directing the weapon(s) at the target. Sensor typ include those that measure target and aircraft motion or position, and those that assist in target detection. Externally mounted sensors should be housed in aerodynamic fairings wherever possible, and protection should be provided against such environmental conditions as handling and accidental ground maintenance damage. Sensors that produce elactro-optical or electronmagneic energy should be located so that neither direct nor reflected energy enters the crew compartment. Furthermore, the mounting provisions for auch semors should permit attachment of ground operation warning devices to alert ground crews to potential radiation hazards. The mechanical interfee between the sensor awd the helicopter should be desgned for adequate strength, ease of maintmanc, and amcurate a:gnmeat with the helicopter datum plane. Sensors projecting from helicopter mold line should be iocated so that they will not interfme with aircw entry and exit. Aerodynamic soeors includc those or the pitot tube, anglof-attack indicator, and air data computer. They should be located as iar forward on the aircraft, and asllffar from the fUseg or appendags ralmqe
AMCP 706-202 aa is iwan"eal to misiimise aeirodyananlc ismrfeseme and loWa flow variations duo to the influence of the maiW rotr(s). Seemo accurocy levids shiould be aehct for compatibility with fire control accuracy
"reuiresomes
-
Sensors bested laternal to the helicopter .bdl be potecled pine the vibration shock mad power load s*gaviomsna associate with the specific bheliopter design. is genral any amunsing equipment located withi the crew congpaalment must be capable of wlthsamdi~fsg crash load factors without dletachingS trwo its mountings. Vibration isolation should be Provided in accordanc with MIL-E-5400 and MILSTD-S 10. power requirements (including number and sin of delectica wires), the necessty for shielding, and/or the use of nonstandard electrical connetor should be considered in the design of th internal Mounting strucure. Mounting provisions for all sensor units should provide for easy removal of fostsiners arad connectors, and for structural dearDs= aidequate to permit rapid removal and/or repair durnn. "'w~ano operatons Scumor using elecro-optical or elecromagnetic energy shoul be located in regions where they will evnm~mm-p minimum electrical interference from otbgr aircraft equipment. This principle applics to both th internal arnd the external (transmitting) portiois, of the equipment. The electrical Power requireiamts of thes equipment types can be significant. In order to minimize transmision losae and elecrical inter~feroene it is necessary to: (1) locate the sensor unit Power supply in close proxi'mity to the balicopter power source, and/or (2) minimize the sepsration between the sensor unit and the point of air cato tavo iid otnti hnazadsition tahe ul bcew during to &voi poeta
u)
aad
otearrwdfn
3-4.5 COWI¶TFVJIS The airborne fire control computet is a specialpurpose device that accepts quantitative informnafion, arranges it, performs a mathematical calculation, and provides qualitative outp~it information. This definition describes a simple Clectri Wl computing circuit as well as a digital comiputer. The specific requirements for the computer are astablished by the degree of fire control accuracy desired. In addition to suppoiting rfie control compuitation, fth airborne computer may be employed to aswist in flight control, navigation, and communication tasks. The specific design requirements of the computer syatm AsW be in accordance with the governing desig requirements of the fire control system for
soaing heicopter nar,gaption-vionlc computation flancion. Power rosjiiianLni AoU be as specified by comnputer design requirements. Computer vibration mclation will be required as spedifod in the deAig mn quirementa In meeting computer aWoes requiremcnts. consideration must be givin to removal. safoty, replvcsment, and component inspection. 3-4.6 FIRE. CONTROL ACCURACY Guns and rockets should have adequate structural support in order to minimize helicopter structural deflactions during firing or launching. Optical sights should be located so as to avoid the sighting aberrations of canopy distortions. Sights should be installed upon rigid mounts, but without inducing undin sight-fine vibrational distortion. Locations of sights and armament should be such as to avoid ti: probability of excessive parallax errors in the fire control computation. The sight and armament subsystems should be installed with suitable adjustment and lock devices for proper boresighting and barmonization. Electronic equipment should be prot £td from the noise generr~ted by helicopter power Pý2ýf
etk
.nu Wu0 har i'.Mw~nonta
.
14-1W#t,
generate dect~rnic noise). Data sensors and ballistic computa- should be chosen so as to provide cornponent accio-wacy characteristics consistent with the sytena amcracy requirwrients of the governing helicopter specification. 94A low"i Stoblratbou Fosmeapito tblztonfthsgt oraisomedapicainsaiiain ordetheoehlcper soight dynamics as a source of sighting error. Stabilization is also incorporated in automatic target tracking equipment. This equipment typically is designed as part of the sighting station, but remote auxiliary comnponeat location may be rquirod. Auxiliary comnponents should be installed in accordance with design practices specified by the sight supplier. 8-4.6.2 FIre Ceotrel Datue Plasie A physical reference surface should be established so as to relate sight and armament equipment for basic alignment and harmonization of the fire control system with the aircraft structure. !n a fixed weapon system installation, the weapon firing line generally is aligned so as to be parallel to the aircraft datum line. The designer should consider weapon characteristics, such as tangential projectile throw and/or barrel cant, during the alignment process. If the system is radat-directod, the radar line must be aligned parallel to aircraft datum.
1
S,
.
"
AWC 70&-202 For flexible fir- control systems
-
either radar- or
optically-directed - provisions shall be made to align and test the line of sight, the tracking line (turret minus weapon), and the wetpon firing line parallel to the aircraft datum line. Provisions must be included for testing and harmonizing the fire control coordinate system wit, the aircraft coordinate system. The precision of measurement is dependent upon the overall fire control system requirements and the associated error allocation. -4.6.3 H In order to provide the greatest firing accuracy, the major components must be adjusted carefully. This is true especially of the sight and its accessories, the projectile-launching equipment, and the aircraft itself. This harmonization involves the orientation of three reference lines in the aiicraft: the aircraft datum line (flight path), the sight line, and the armament line. Lugs aligned with the aircraft longitudinal axis should be installed to provide a surface for leveling the equipment and for establishing reference lines. Harmonization of fixed weapon installations generally can be accomplished by mechanical ad,_st.m.lnt of the weannpn. For flexible installations. mechanical and electrical adjustments are required in order to assure coincidence and alignment of sight line and weapon line. Harmonization techniques can be parallel, point, or pattern; and the technique best suited to the specific weapon type should be selected. In parallel harmonizotion all armaments, plus the sight line, are aligned parallel to the Armament datum plane. Point harmonization typically is used when the armament is installed well outboard fronm the centerline of the aircraft or the sight location. In this case, rall arii-,anit i-s-" ...... S .•ht, 1 c. .;.g I;. rtor_ sect at a point ahead of the aircraft, thus con€cntrating the fire of the weapons upon a single, small area at a selected range. Pattern harmonization is similar to parallel harmonization, except that the parallel armament line is elevated above the sight line by gravity drop or velocity hump corrections. All use of a special n techniques harmonizatiz taroet( temonizatien require board) the as supporf equip-
tent. ment. COMPONENT LOCATION The major points to be considered in fire control component installation include vibration and shock isolation, cooling and heating, radio noise interference, accessibility, electrical shock hazards, and crash safety. All weapon system components should 8-4.7
be mounte¢ so that the entire travel of the shock mounts is possible in all directions without inter-
8-10
ference between components. For small cornponents, the shock mounts should be adequate to support the weight of both the component and its cable connections. The components xh/l be located wherc they will recive an ample supply of circulating air, particularly when airborne. Heating may be necessary for some components of the fire control system ander extremely lowtemperature conditions. Radio noise interference should be rhduced by grounding the case of all zomponents securely and filtering power supplies, and by using shielded cables for pulse-carrying applications. Electrical shock hazards will be reduced greatly by secure grounding of all components. Ease of installation, alignment, and trouble-. shooting should be considered during equipment design so that connectors can be disconnected readily, even under adverse conditions. Components should not block acoess to other components. S'rviceability of equipment should be enhanced by locating adjustments and test points on a single, accessible surface. Where this is impossible, the use of slideout racks to permi: removal of equipment from shodk mosnts should be considered. Ii the cornponeni must be removed from the aircraft for adjustment, sufficient cable length should be available to allow the component to be removed and placed on a service rack without disconnecting its cables. Great care should be taken to insure that avionic subsystem components cannot enter crew spa= as lethal missiles in the event of a crash. Crashworthy component tiedown strength and/or crashworthy barriers should be provijed in order to overcome the lethal potential of avionic components. A-5 ANTENNAS ".11 GENERAL The communication and navigation equipment d'scussed in pars. 8-2 and 8-3 requires a variety of an. tennas, ranging from those in the low-to-microwave frequency spectrum to horizontalradia. and vertical polarization andthose thosehaving with different vetclparzioanthswthdfrntai. tion patterns. Antennas are susceptible to rotor-
induced modulation, triboelectric cl,..ging, and noise generated by corona discharge. The combination of electrical requircmcnts and problems created by the platform presents the antenna designer with difficult design requirements, Safety problems are important, It is essential that antennas capable of emitting po. tentially harmful or fatal radiations be marked with appropriate warning labels to proecude fatal or serious injury to aircrcw or maintenance personnel
during normal operation or ioutine maintenance.
(
AMC
S
41•
ANTENNA DEVELOPMENT
The development of a helicopter antenna -abased gpon the requirements of the associated equipment and the mission for which the helicopter is intended. Some of the con-iderations are: I. Frequency range 2. Radiation characteristics 3. Polarization 4. Efficiency 5. Voltage standinj,, wave ta-z.o (VSWR) 6. Noise 7. Environment 8. Structure. The frequency range of the amtenna determines its basic dimensions. Army helicopters use frequencies from 150 kHz up to the visible range. Without conridcration of power output, the high-frequency region extends to 30 MHz. Depending upon the time of day, the time of year, the sunspot cycle, and the vagaries of the ionoaphere and its various layer3, the HF band is considered to be the best for longdistance communication. With L
)
0
a
a
(SSB) equipment, high-frequency communication is becoming more prevalent. Howeyer, mission requirements will tstablish urage criteria, For VHF, 30-300 MHz and up is considered to be line-of-sight communication or propagation, depending upon radiated power and receiver sensitivity. and upon the points of radiation and r%ception. For example, a reasonably clear area provides true line-of-sight communication, while multipie layers of vegetation, such as are encountered in Southeast Asia. require much more power in order to effect through-vegetation transmission of electro-
-
-/ \.
magnetic impulses. For operation at frequencies in the LF and HF ranges antennas are quite long--eg., one-quarter wavelength at I MHz is 246 ft. Because the dimensions are so large, it is standard practice to have the antenna system include an antenna tuner as a coupler. Ing coupler automatically matches the impedance of the electrically short antenna to that of the transmission line. This method of loading wire antennas becomes leas efficient as the ratio of antenna length to the wavelength of operation becomes smaller. For frequencies at or above VHF, the sizu of the antenna is less of a problem. As the ciectrical length increases, the instantaneous bandwidth of the antenna also increases, and the result is operation over a wider bandwidth withoe't tuning. Whercas the wire antenna must be tuned each time the frequncy is
S~8-1l
706-202
changed, the VHF and UHF antennas are fixedtrned, and are capable of efficient operation, with low VSWR, over a band of frequencies. Decause of the wide instantaneous bandwidth, the antenna also can be used simultaneously by different equipments tuncd to different frequencies. Diplexers and hybrid devices are used to provide ivolotion between equipments using the same antenna. Radiation patternb of the antenna indicate where energy is being radiated, or, conversely, from which direction it can be received. Communication and direction-finding equipment generally requires otnnidirectional radiation in the azimuthal plane, with the maximum amount on the horizon in the vertical plane. Navigational equipment requires radiation in specific directions. (Because of the physical geometry of the airframe, truly omnidirectional patterns never are obtained.) The airframe directly influences the radiation by its shadowing and re-radiation effects. The airframe can radiate energy coupled to it at frequencies where its dimensions are an appreciab!, part of a wavelength. At higher frequencies, the airframe blocks and shadnww raddiatinn in evtain direttinnq
The relative positions of the antenna and the rotor also affect the radiation patterns. The effect of the rotor is to modulate the radiation pattern at a frequency determined by the number of blade passages per second over the antenna location. The carrier frequency, along with each sideband, will be modulated by this frequency. As discussed in par. 8-1.4, this modulation interferes with the performance of equipment that makes ust of information contained in modulation components close to the same frequency. Rotor passage near an antenna also can affect the -am. of"the..c c" , whI.... r• "tt- in.... a "m-"h l t signal. From empirical data, it appears that an antenna can have a peak-to-valley variation of about 6 dB and sharp nulls of 30 dB without cxperiencing overall degradation of performance. For the sake of economy and practicality, it is imperative that the best antenna possible be provided. Mission requirements will determine the selection and use of radio type(s) and associated antennas. For instance, the VOR pattern is optimum in a forward direction, while tactical communications, IFF, and other primary radio aids should be as omnidirectional as practicable. The polarization of the helicopter antenna must correspond to that of the antenna at the other end of the communication link. This requirement does not apply to HF aotennas because of rotation of polarization by the ionosphere. Cross-polarize' signals can be radiated from linear antenna elements as a result
.
of reflections and currents in the airframe. The crows Measurement of impedance and mutual im-K., polarized component represents an inefficiency or a pedanoc is accomplished most easily on tho belipower loss, and is a consideration when selecting the copter although a full-sc~ale mock-up, containing antenna lcamtion. those portions of the helicopter within several wave,'The effocts of tribotlectric (friction) charging inlengths of the antatnna, gives accurate results. AD fluence the selection of antenna location. The helimodel measurements must be verified on full-scale copter airframec and nearby antennas are cha- god aircraft. This technique is analogous to aeodynamicelectrostatically by the rotor downwash, and voltagat model testing in &wind tunnel. high enough to produce corona can result. The corThe present range requiruanent for VHF-FM cornona will occur on ahaip points. or on points of high munication having a 10-W power output to a nonelectrical stress concentration. If these arcas are on matching antenna is 40 mi. This is a reasonable rang antennas, oi are electromagnetically coupled to when tested in the optimum condition. antennas, the broadband noise generatod by the corWhen an antenna is to be used over a wide firequenon: will be introduced into die receiver. cy rangt, i.e., VHF-FM 30-76 Mliz, it must he Antenna location must insure that die antennas are broadband if efficient operation is to result. If the decoupled from each other; this is practicable for antransmitter/receiver mismatch is not too great,. U tennits operating in the same frequency range. reasonable efficiency will result. VSWR, the Voltage -ý Mutual coupling between antennas affects-their imStanding Wave Ratio, it the criterion for acceptance pedance and the radiation patterns of individual eleof antenna matching; it is determnined by the ratio of ments. High voltages can be introduced in passi-ic cirforward to reflected power. The Radio Technkca cuits if excessive mustual coupling exiists between Commission for Aeronautics (RTCA) aW chowe a transmitting and receiving antennas. VSWR of 5:1 as an applicable standard. However, a Dcvelcpment of helicopter antennas depcnds upon more stringent ratio is required in many came. Comnr~ne oA~. ,~trn i anta.,,na .hst.,,L4 hows-~ a VQlWA Aff I A-1 th.~ ~ ~~yp.,.,.,,.a ~ ~ ~ ~~~~~~.ne MIL-A-25730. When employed with discretion, the o. less. TACAN and trnusponder antennas should usc of scale models for antenna development pro. have a VSWR of 2.3:1 or less. VSWR is okcompla. ( vides the ability to predict the suitability ot unsuitratio of incident powtv to reflected power as given by'ability of an antenna location. Usually, the model Eq. 8-I. technique is used on lar~ge prototype aircraft with the I Reflcted Powur model scald down to 1/20th to facilitate manipuForward wr lation. The scale factor is variable, and any scale facVSWR :t 5 10c tor maky be used provided that the resultant freiv*~os ored o~ quencies (which mumt be "ca.idupward by the sain scale factor) are easily obtainablu. ".3j IOA711rON AND INSTALLATION OF' Frequencies in the Sgigahertz regpion are not easily ANTENNAS scale. In addition, frequencies in the HF region In most casta, thot mveral types of antennas often are not scaled. On a 1/20-scale model at 15 that can be used for any gliven item of communication MHz, for example, tht; model frequency must be 30 or navigational equipment. and the final choice Is deMHz and thesw two frequencies inherently propapendent upon the requirements of the specifkicngate differently. Theory is quite logical for scale stallation. For additional information governing anmodels, but care must be taken to scale every detail tenna installation and location, see MIL-STD477. affecting; antenna performance. For example, caling During the initial design phase, incorporation of wire size in order to maintain identical current dissystem zero-drag or flush-mounted antennas must be tributions must be considered. The model must be considered. It is extremely important to submerge the isolated from its surroundings, and consideration antennats, not only for increased flight efficiency, but must be given to ground reflection& and to radiation also to minimize maintenanov problems. from connecting cables. In most case space limitations prevent the location The model measurcineats ame made in the scale of antennas far away from~ one another. dimensions and for orthogonal polarization. The Specifications usually require a three-eighiths waveradiation patterns arm measured over the scaled frelength separation, but this ir unrealistic; therefore quency range of operation, for varying positions of separation to the greatest de~.me possible should be the rotor and for different antenna locations. The opmade. An alternative solution is the multi;1sage Of a timumn location of the antenna with respect to radiasingle antenna, with passive devices added to aid in tion patterns will result from these measurements. multiplexing. An excellent example of this usage is a 8-12
.~~~
, ., .., .........
.•,•
C•706-M,0
ipr-dra& wide-band antenna for both UHF and VHF freauencis. With the assitance of a diplexer this single, simple antmna is used to receive and transmit simultaneously in tha UHF and VHF communication bands. The same philosophy could be used for lower frequency VHF (30-75) and ADF sen antennas. W-5..1
I
,
Cemmutmscad Astema " CeulkMuajem Together with EMI/EMC, .considerations, ontennas poe the most difficult problem for the helicopter avionic engine"r. A typical helicopter with standard communication antennas is shown in Fig. 8. 3, which depict a simple, operational combat scout or observation helicopter (none of the navigational &atemnaw©s e shown). Antenna functions are affected by antenna licatice, and the helicopter in Fig. 8-3 illustratca typical problmsm For exumple, due to the HF antenna 1ocadon a hard landing could affect operational characteistics. This antenna also is beaten by the "aim rotor downwash, and breakage could caus the ansenn to bccrc wrapped up it. the tWil rotor. In addidioe, an eclctical impedance problem results from the main rotor blades passing over the antenna and causing rotor modulation. The VHF-FM #1 antenua, used for tactical communicationk ist in the exhaust, which could camu.physical degradation, and the bulk of the helicopter is forward of the antenna, resulting in partial antenna shadowing. The VHF antenna, shown in the belly, may be relatively clear of many problems, but the landing Sear would gave reflective properties and consequent nulls in the antenna pattern. The UHF antenna, shown forward and above the cab top, is vulnerable to triboelactric noise and to rotor modulation. The VHF-FM #2 antenna is in front of the aircraft and in the field of view of the flight crew, which could be distracting. It also may be vulnerable to rotor modulation and triboel-ctric noise.
-,
S.... so Aý
.Radiation
.2
N
Stransmitting Fig
8-3. Typical Commmmleatlom An a Layoet
Sdirectional,
Effects of rotor modulation can be controlled offoctively by installing notch filter equipment in the primary area of interest. In the case of communication receivers, band stop filters are used to eliminate unwanted modulation frequencie in the audio band. "-3.2 Low Fre"q (LF) The primary use or the low-frequency sp%-trum is for automatic direction-finding. The ADF system use" a loop antenna having a figum-of-light radiation pattern, plus an omnidirectional whip (sese) antenna. The sense antenna output is combined with that of the loop antenna to produce a cardioid pattern, thereby eliminating the directional ambilguity of the loop. The location of the loop antenna is restricted by two considerations: 1. The cable between the loop and the receiver input is part of the receiver input circuitry and is of fixed length. 2. The loop must be located in a position of minimum pattern distortion. The magnetic field lines that induce a current in the loop are distorted by the airframe, thereby causing an ,,,, ,,v.. ,,,, compensated in the equipment, the design engineer must determine the best location for each installation. The sense antenna should be positioned in an aca of minimum electrical field distortion to maintain accurate ADF performance as the heli. copter flies over or near the ground station in what is called the "'confusion" zone. The size of the confusion zone, in which the ADF indication can vary as much as 180 dog, depends upon the characteristics of the sense antenna apd upon maintain'nS a minimum signal input level to the receiver. 3-5.3.3 High Freqnracy (HF) HF, employe d for long-range communication, use wire antennas. The wire can be fixed between two points on the helicopter, or a trailing wire can be used. The use of antenna couplers is required with this type of antenna. A major problem with wire ante,%nas is the posabdlity that they will become tangled in the rcAor-. patterns, which, ideally, would be omniarc dependent upon location. They can be shadowed by the airframe, which itself can radiate and cause distortion. The wire antenna usually will have nulls at its end directions. The antenna wire is coated with polyethylene in order to prevent corona, and the supports must be designed to withstand level voltages. LORAN also utilizes the wire antenna. Corona and voltage breakdowns are not problems, but omnidirectional co'y.agt still is a requirement. 8-13
8.-.3.4 Very H1g0 Frqme~y (VHF) The marker beacon receiver operates at 75 MlIz. The antenna radiation pattern must be downwardlooking, and must be polarized parallel to the axis of the helicopter. There are several antennas that will meet the radiation requirements. One is a balanced antenna mounted under the airframe. Other suitable designs are loaded half loops and flush-mounted cavity elements mounted in the same location. The glidescope receiver operates in the frequ:ncy range of 329-335 MHz. The glidescope antenna must be designed for reception of horizontal polarized signals with minimum reception of vertically polarized signals. The antenna must be located forward on the aircraft for proper glideslope reception. The VOR operates in the frequcncy range of 112118 MHz and requires a horizontally polarized, omnidirectional antenna. The ILS receiver operates in the frequency range of 108-112 MHz and uses the same antenna as does the VOR. The location of the VOR antenna is critical due to rotor modulation effects. The VOR determines the phase difference between two received signals modulated with 30 Hz, which is about the third harmonic of rotor-induced modlation; if irttilfetreijc iu too gSrat, tht VOR is inoperable. A loop mounted on the undcrside of the airframe and configured with a vertical axis is suitable for the VOR antenna. Location of any ansenna on the underside can result in airframe shadowing in some direction. The rem's horn (a modified dipole), stacked dipoles, folded dipoles, and Pushmounted cavities also aJ! are suitable designs. Communication equipment operates in the frequency ranges of 30-76 MHz and 118-150 MHz. Some form of monopole is used most often for communication. The position of these antennas is determined by the necessity of obtaining omnidirectional radiation patterns. The antennas must be designed to prevent corona discharge, and shculd be decoupled from triboelectric discharges that would introduce noise into the receiver. The difficulty of ob-
I
8-14
taning omnidirectional coverage sometimes can be overcome by using two anennas. If the radiation patterns of the two antennas are complomontasy and the antennas arc isolated from each other, they can be driven in parallel with appropriate impedence matching. "-3.5 Ultra High Foo.SCy (UHF) UHF communications operate in the frequency :ange of 225.400 MHz. Monopokl-derived configuratiods are used as both UHF and flush-mounted types, such as the annular slot. The gmeral requirements for VHF location and intallation pertain to UHF as well. 8-53.6 Special Puopue There are other types of equipment that require antennas, but these are limited in use. The considcrations for location and installation oC' these antennas depend upon individual system requirements. Doppler radar antennas, for example, always are mounted flat in the bottom of the aircraft. In general, the same restrictions discussed heretofore apply to all types of antennas. REFERENICES I. USAECOM Manual, Interfereuce Reduction Guide for Dejign fgineers, Vols. I and I!. AD 619 666. 2. AMCP 706-125, Engineering Design Handbook, Electrical Wire and Cable. 3. John G. Mast, Efimintaon of VO.R Sixgl Rotor Modulation, Paper 924, 31st Annual National Forum of the Ameriam Helicopter Society. May 1955. 4. Myron Kayton and Walter Fry, Avionics Navigalion Systems, John Wiley and Sons, NY, 1969. 5. Turnbow et ni., Cra'hSurvival Design Guide. TR 71-22, USAAMRDL Fort Eustis, VA. Revised October 1971.
V4
706202
____ __-AMCP
CHAPTER 9
HYDRAULIC AND PNEUMATIC SUBSYSTEMS DESIGN 9-0
LIST OF SYMBOLS
g H lf. dH/dt
-
-
acceleration due to gravity. ftAWc true prcssu~re altitude, ft measured pressure altitude, ft change in pri-sure altitude with time utt
standard sea level conditions, ft/se AM - height difference, ft A ,- altimeter position error correction, ft P -static pressUre. psi P, measured static pressure, psi P, - total pressure, psi P. - measured pitot pressvre. psi AP - prusure difference, ps P - true atmospheric pressure, psi 0 - true im~pact pressure., psi - mcusuire impact pressure, psi -true
A Y, PSI
airsneed. ninh
- -..
calibrated ai~pýkt
-
measurcA calibrated airspeed, kt airspied pow~tion error co:rection, kt dcnsi!.y at standard sea level conditions,
slug/!ft'
9-I DUCIONplacement NTR 9-1 ITRODU TIONnominal Chapter 9. AMC? 706-201. describes the many design trade-offs neoces~" in the final selection of secondary power subsystems. This chapter deals with
~~,~tjEk.
~
H
~
RA
M
ICSIS'T
9-2.1 FLIGHT~ CONTROL POWVER SYS9l?.OvS iclicopter ffight control syitemý may vary in cornpluxity from the racativcly simple power boost system with manual reversion to multiredundant systems wht~re each system is designe-d to provide the full power requiren icj operalc tat flight contrvl funictions throughout the vchi-lc performance env'elope. The nmultirediiniant sysf-:n is discussctd in this chapic! because it contahib~ the basic elzments of all typcs of systems. This typr system s-haII be e-mrloyed uniles ihe aircraf, can be ct~atrolle.-i without boost. A
-
*j.
tions as engine starting, auxiliary utility systems, and emergency backups. Pnicumnatic power also m'ay be usvd fur auxili&rý puw.-r unit (APU) starting. 92 H D A LCS M TM
~
.
,..Iht--
.
cImm
Presented in Fi$. 9-1. The system shown contains its own fluid power ge-neration. fluid transmission, and fluid suppuly components.
9.2.1.1 Central Kydraulic System' Hydraulic power is gene~ated by variable-dispumps that ate comlpensatcd for a system design pircssurc. Fluid may be supplied to the pumps from &u-pre~ssuri&u% c. bastrap reservoirs. The pumps vre driven by an accesory gcirbox, which, in turn, is driv~i bý the transmission whrn t'nr roioi(sl
IArfing kiS and by an
systems. Hydraulic applications primarily include flight
control and utility functiotis. Flight control functions include servo control of cyclic pitch, collective
VENT
pitch, and directional surfaces. Utility functions may include part or all of the following: 1. Personnel/cargo hoists
CUL~.A1fT 1D
PPCEI
2. Cargo hooks
3. Loading ramps
R*
14
L
4. Doors 5. Landing gear 6. Gun turrets and drives
)
7. Rotor braking 8. Wheel braking and sleering 9. Engine starting
-WR PLOPL
LE
CDISCTIM
10. Fluid dampers.
Pneumatic applications, while not as widely used in helicopters an are hydraulics, may include such func-
FIgure 9.1). Central Hydraulic Sysiemn 9-1
AMCP 706-202
Ati~ns,
I
*
APU during ground checkout. Ccrntral system componenrts include: 1. System manifold, a package containing system Ifiiters, presure transmitters (with associated snubber and fuse), system relief valves, ground tesi connec6nd return line chtxk valves 2. System ac~cumulator, a gas. -. echarged, piston unit with associated servicing valve ard pressure gage 3. System reseryo-r, including return and pump suction line fittings, bleed aild fill pravisions. overboard vent, and level indicator. The reservoir also incorporates reservoir level-scrnsing, with associated subs:'stun isolation valves,
The stick boost system provides the pilot with lowforce stick movement capability, As shown in Fig. -94, an actuator is provided in each axis to overcome friction and inertia loads. Thr. actuators function to react inputh from the GAS so that they cannot be felt through the pilot controls.
9-21.2 Flight Control Sotseystes
samec components
A typical flight control subsystem consists of (1) a boost-actuating system, (2) a stability augmenitation system (SAS). and 10) a stick boost hydraulic system. The pilot's control movements. transmitted through a system of bell cranks, rods, and levers, arc mixed to provide the correct lateral, cyclic, and pitch motions through hydraulicaly powered actuators as shown in Fig. 9-2, If dual actuators are used, each half ol .the se'aacor is porwertid by a se"rate system. A typical duJ reversed SAS i:: ii.ustrated in F.jg 9_ AS actuator inputs to thec boost actuatorsC 3~. Th Sc affect the morement of the rotor bladC5 Without feedback forces to the pilot controlb. The SiAS actuators
sytmd anduigpklodhseiseaivy high, additiowtl pu~mps in parallel may De neosaiary. T iiietcwih n ieo usse rut that do not have high pressure and flow requiremepesuerdcs hldbcosded 9221EesSI~Sby~u 9221Egn atn usse There are two basic types of hydraulic enginestarting systems. One uses a limitod amount of stored encrgy that is available in an accumulator, while the
mist be capable of being engajed or disengaged by
tho design goal is to completc the start inthe shortest
9-2.2 UTILITY HYDRAULIC SYSTEMIS Teuiiyrse a epwrdi setal a epwrdi setal h tlt ytr the came nmanner as is the flight control system, using accessory gearbox-drk-n pumps when the rotor it turning or the APU for ground operation. The central portion of the system will contain basically the
as does the rlight control system. If
u
puc PUiVG L1n'.,jiiuaibauzb 11MII41
Au VWLUa
from an auxiliary power supply. Becausc energy is imitcd in tlie first type ofsystcin,
feature that oper-,tes in the event of hydraulic power___________TUI
SUM
faiure.
~TFTTDUAL
j.
Figure 9-2. Dual System Hydamlic-powercil FIgNt Controll Actuators
ACTUATOR
Fewue 9-4. Duai-p#.wervid Stick Dow
Hydrumak System wL.L W O
-I~-' - - OVLVE
.S*
Figwrt, 9-0. [)ual-poweyed Stablity Ampmawathiu Systew
£MN
1?T
WALK
60"tIm
Flswe 9-5ý Hydrainic Staritms; Ean.gy-limbiu Systems
-pump
*ccomplished by r~claing stored energy from a charged accumulator. This eaeirgy di ives a positivedispiacemcrit starter mounted on the engine. The starter may consist of a simrlc fixed-displacement Motor, and is sized so that no special coatrols art needed. Weight can be minimized by using tht: highest pressure and largest displacement acceptable to the motor. These two factors are limited by the torque capacity of the mounting pad and the fuel control acceleration capabilities of the engine. This tvpe of starting system is most appropriate foi- star-ting small enigines (50-150 hp), where the accumulator size and charging time are not excessive, The power-limited type of starter usually consists of a selfaufficient styatem that uses a small turbine en~gine as an APU (Fig. 9-6). Hydraulic fluid isstored in an accumulator that has been pressurized either by hand pomnp or by a previous operation of the hydraulic system. The pressure released by energizing a solenoid-operated starter valve drives a fixed-displacement starter pump, which acts as a motor to start the APU. When the APU starts, the starter converts to a pumping mode so as to drive the matin ..ngint. starter. This starter ag'ises the proper flow or pressure, and its displacement is vared auto-
)
!n~waII,, ti anatn ve~6roto-t
the. pni,
np tdI ;Alrwm-A
The power limit of the APU is not exceeded because ACCUMLATORActuation AIWLATOrelease
GAS PU" LUfeature CWC
of this variable-dispincenent feature. In powerlimited APU systemfs, the starter/pump is tized (or the maximum output capability of the auxiliary engine because the starhirf requirements are lower than arc the pumping requirements. The system shown in Fig. 9-7 uses an energy. limited, dual purpose. starter-pump system on the APU, and a power-limiled, variable-displacemert starter on the main engine. As the APU isstarted, the stanter/pump drives a fixed-displacement motor mounted on the accessory gearbox. The gearbox motor drives all accessries, including the utility pump(s), which in turn provides the power to drive the main engine starters. After the &tartcycle iscornpleted, the main engine drives the accessory gearbox. An added advantage of using an APU starting system is that it can be operated to provide power for ground checkout. 9-2.2.2 Cargo Door and Ramp System, Cargo and/or troop carrying helicopters normally will incorporate some typc of cargo door and ramp system. The system shown in Fig. 9-8 is actuated by two direct-acting hydraulic cylinders. It is important to note that the actuators are self-locking in the rptrnicl'IM nn~titnn The m~animI enntrnl valus-
of the control valve directs pressure to
the actuator locks, and the ramp then ispulled down by the force of gravity. This is an important because the actuators may be unlocked via hand-pump pressure when utility system pressure is
MOTOR
UNPI-P
T 0 'RAMP
TURIMAPUI
RAMPCONTROLVALVE
)cT
HydrmI&t: S~amilig,
O~p,
RA P UP
O,
STO
E R TURNeI.
~fs
Pmw-lifted System
AV
ACCWULATRT
11AND PUMPCA$ CFEL-
O
ILUID
V~ALV
TO
SV
T
ItsSTAR
AGTORF
FKIN 9-7.f APUM IatqSse
k~p
avai~able, allowing the ramp and door to free-fall as an emergency measure.
STARTSIMInot VALVEVALVEopen
Flpre 9-.
hanll
located conveniently near or adjacent to the ramp.
RA
lr STA9-3
4
sg
MPCO IA
oradRm
R
AMD
yn
A4P 706-202 In normal operation, initial movecment of the r,,,mp actuawes a sequtrnce valve that perm-t~s flow to a constant-displacement motor. The motor may be moumcd insidge the ramp, and connected to the cargo door by~ avi endless chin, so that it can be retraciod into the ramp structure. The sequence valve blocks downstroke flow from the ramp actuators until the door is completely recracted. When the door is positio~ned pro, irly, a hydromechanical stop halts flow through the motor. Decay of .'-otor back pressure Opes the sequence valve, allowing the xasp actuator to bottom out fully or to travel to the point where the. ramrp toucher, the ground. Closing of the ramp is essent~aly the smain but in the reverse order. 9-2.2-3 Cargo aud Piersomad FIlst A utility hoist can be provided for loading and uni loading cargo and for rescue opierations. The hoist, as shown in Fig. 9-9, is powered by a hydraulic motor. The motor requirements are established so as to provide a particular hoist weight capacity and maximum reel-in speed. The speed can he rmade infinitely variable within the rated speed range by means or a iiydraulic sei-vo val-te. The ser7vo valvtconrtroll signal if. generated by a potentiometer incorporated into a control knob in the cockpit or the hoist operating sta.tion. For the hoisting optration, pressure is directed to the "in" port of ihe motor. For extending the hoist, a pressure reducer should be used to provide the relatively lower motor torque needed for cable extension. A flow regulator incorporated into the retuin line, downstream of the control valve, regulates flow in both directions. Limit switches can
HiOIT "ARAE AELEASE VALVEE
L-w~ovided to stop the hoist motor at the full roed-in ant; full reel-out positions. A.hydraulically released, spring-loaded. "on" brake can be provided for failsafe operation of the rescue hoist. Provisionts should be incorporated to stop fthmotor as the hook pasaw through the cargo-rescue hatch in order to preVqnt whipping of the cable. In addiion, a device can be incorporatod to keep the cable under tension under noload conditions. 9-21A4 Rotor Brake The rotor brake shtl be capable of stopping t?.e rotor within a preselcrted tinme period for a specific range: of rotor speeds following iimgine shutdowi., and the brake 4Ad1 be cap.;ble of holding the rotor stationary when tull ground idle engne torque is applied. As shown in Fig. 9-10. the biake may be applied by a sokaoid-energized valve that directs pressure to the rotor brake. A precharged accumulator provides a steady hydraulic pressure for braking whtn the main system is depresssrized. A presiare switch should be incorporated in order tr' provide a cockpit warflnla lirht indacaLion when. th" b.3hc is on. To rees h rkthe soeodvalve is deenergized to the off position. allowing the brakek. pressure to bleed off to the return system. Springs may be, used in the brake assmbly so as to overcome the return bystemn back pressure in order to allow separation of the pocks F~om the rotor disks. The rotor brake skall be desigrned to be fail-safe. For sL.Wal helicopters a simple rotor brake system may be satifactoly.
RTLB
-
E VI -E1KRriLW
BRAKEON
C,
ROTOR
AAR
HOISTCONTROL I' WALVFASSEMBLY
FLO
N R LC ULA ' DO L Nh R TEt TO IT~ E.YAS REE[ OUT-.....,..........R
1 U
VIBISCH
CKLCK VATI'
ACCLIV,1A TOR
SLIUT
0, OECERLRGIZE
8RAAE
CHECK (WYILKA
'KTOR'h MOISTIIytKAUL;C
ROTORDRAKE VALVF
F11pe 99. arpLLAPeimaiand Hi~st
Figure 9-d ~U
rnmourte symto
Figur 9-1a it""s eeke Syas"
9-214 Wheel Brakes Hydraulically powered wheel brakes srWl be designed in accordance with MIL-P-8585. The circuit shown in Fig. 9-t I represents a braking system with differential control. Each brake pedal provides a direct input to a master brake cylindt . Hydraulic pressure may bc reduced at necessary for brake operation. Parkinig brake capability also should be incorporated. A warning indicator shall be provided to signify when the parking biakes are "ON". 9-2.3 HYDRAULIC SYSTEM RELIABILITY Good hydraulic system reliability cani be obtained best by recognizing the probable weik links during the initial design. Because the system must be designed within the constraints of weight and cot and must meet appropriate fIilitary Specifications, reliability aspects of the design must be optimized.
*
)
9-2.3.1 MWgh CostOe Redowlasey The most critical hydraulic failures are those that ultimately cause los of the capability to operate the prmnary flight oontrols. MIL-H-5440 requires that. whcrvtý hydraulic powcr is u=-d for the primary ~flight controls, a completely s"prate system s/sol be provided for that purpose. Further, it is required that if direct mechanical contro' is not suffickiet to allow controllability as defined in M IL-F-8785 in event of a hydraulic failure, an emergency power source shalil be provided in ordc- to supply the kwne4esry controllability. As.a. cmu of mieeting the redundancy requiremerits, several design techniques may be considered: 1. A an gi primary Right control system with medianic-al reversion. This method is simple and relatively lightweight, and can be used if opeiating lout are bot above pilot and/or structural capabilities.
TO,9-23.3 U
TIO.
Flpw 9.11. Wheel Dbab Syssem
2. A single primary flight control system with integrated, electrically powered hydraulic backup. The survivability characteristics of this design are good but may result in a heavier and costly installation. It also may present beat rejection problems. 3. Dual flight controls. This may be the best approach if high powtr must be: delivered to the flight control actuators. Dual or tandem actuators can be used to enhance reliability further. However, this type of system involves more wcight, more components, arnd, therefore. higher cost than other -approaches. 4. A single flight control system with active utility system backup. This is an approach which should be considered if the utility system can supply power sufficient for the normal utility functions. as well as approximately two-thirds of the hinge mnoment requi.-od for the flight control functions. Use of priop ity valves in the utility system should be considered in order to insure that priority is given to the flight controls in the event of loss of the independent flight control hydraulic system. 9-2.3.2 Utility System Redimilaacy In Sccr~ieas, utility functions are not crAical individuaily to the 'control of the helicopter; therefore., total system redundancy need not be considered. However. specific functions within the utility systeir. may be critical during emergency conditions. If so, emergency or alternate modes of operation should be considered. Usually, the least reliable component within the system is the hydraulic pump, due to its relative complexity, high operating loads, and con. tinuous operation. The use of two u~tility system pumps can provide add'tional reliability. Each pump can be sired to provid one-half of tin. maximum system power demand, thereby providing a good weight and cost trade-off in comrarison with asingle. large system pump. If one of the two pumps fails, systan performance is reduced, but only the failure o( both pumps can cause total system los. Mlaesimemu Relia~lely A*spet The key to high systemn reliability is reduction of the effects of single-point failures or elifflniation of their cause. Major types of recurring failures are: 1. External component leaks, causing loss of system flstid 2. Leakage of precharVed accumulator gas into the
hydraulic fluid, causing pump cavitation or dumping
of fluid overboard. The effects a.^ component Ickap can be alWe viated significantly by use of a leakage isolation device in each flight control syutar brands circuit. The various types of sudh device wre discumied in par. 9-4.5.4. AMCP 706-201. 9.5
z
-.
Loss of systems due to accumuliator prechlarge gas
4. Deflection of components must not cause mal-
leakage can be reduced to a gmret extznt by elimi-
function. Deflection, rather than strength, is often the
nating the need for a main system vccumnuistor. One method is to design a system relief valve having the response capability nwxessary to dump high-pressure fluid to the low-pressure return side of the system quickly enough to reduce the amplitude of pressure spikes* If a variable-delivcry, premsir-compensated pump is used, pressure surge can be reduced by designing the pump with a &rivativ: compensator. The compesator anticipates pressure surges by snsing the rate of resure rise, and thus acts to reduce the magnituoc of that rise bcform it reache a critical value. The design approaches discussed in this paragraph rqeprnt the Lattst state-3f-the-a.n developments and, therefore, may not prove to be the bat metlhod. However, their considerd tAsage could be worthwhile in inceasing system rrliability.
major criterion for design of hydraulic system componcnts and associated structural elements, When this is the situation, the design hall be based upon Unmit loads rather than ultimate loads. 5. Temperature variations must not cause usalfunctioning or excessive stres. Consideratioia must be given to expected temperature variations so that no binding. sticking, or malfunctioning of components will result. Internal stresses, such as those resulting from the usc of dissimilar materials in combination, should not exceed allowable stresse, uns the me4t adverse temperature conditions. Where componaits are expected to operate at extremely high tcm acatuwe, allowable unit stresm may be redred. Hydraulic system design pressures (operating. proof test, and burnt) stall be determined in accordance with Table I of MIL-H-5W40. The design should be based upon the most critical condnion. In Haddition, MIL-H-5440 acquires that all hydraulic systems and components that arm subjected, during operation Of the aircraft, to structural or uFdirhiF '04 not of hydraulic orign shell withstand such load when they arc applied simultaneously with appropriase proof puesurwc as specified in Table I, without exceeding the yield point at the nmximum opera6mg temperature. MIL-H-5440 also requires that actssting cylinders and oth• components, and their a•taching lines aad fittings, if subject to accerated
-2.4
j
HYDRmAULIC SYSTEM STRENGTH CON-,..... . Hydraulic system components and attaching linkages; iU/ be designed so as to meet the most critial loads or combination of loads. Load f-tors or design factors s.a/1 be established for systems in order to insure adequate safety and lift of the components. Where applicable, load factor4 must comply with Military Spzcifhcation nquimrnents. When no Military Spcifiadon requirement exists, these factons must be determined a.Id assigned in acordanwe Swith good den practice. The following basic criteria A lib cosisred in establishing thes factors: i , Tine structure must not suffer fatigue wit., sub"jectedto normal working Ioads. The unit sartis under normal working loads must be limited so as not to eaced the fatigue strength of the material, under repeated loading. for the anticipated ife of the st ture, with str'ws concentration fators taken into consideration.
2. TIh structure mus not yield wbhn sujce to
loads. shl be deiwn and tested on the basis of a prnsure aual to fth maximum pressure tat will be developed, without uceeding the yield point at the nmaximuwms Opatn 00 4ljrj*.
92-"
HYDEiAIJUC SYSTEM TEMPERATURE CONSiDERATIONS Hydraulic fluid selection criteria include th expected mrn of operating temperatures. tiuc at eatreu tamperaturns ompan to available means
of apnau un control, and fluid physial propenris
maximum expected loads. The unit misus at the limit load must not exceed the yield strength of the
at expected temperature levels. Where ambiet and structual tempratures are above the hydraulic fluid
material. It should be reowgnized that test loads may
flashi and/or fire points in a compartment. the pa-
be imposed upon the structum that may exceed the maximum limit load encountmed afher insallation.
te•tial fire hazard iast be considered. Flua sability is affected by thenrma stress, which can result in changes, in visAosity and fonnasion of volatile coaposents, i,-soluble mteriaL, and corroive dslosits. Hydraulic system eff'iencies a•r rodtud by sugh fluid viscosity at lower tampraturm, which rfutam in inlet problems with pumps, sluggs"h repons of critical actuators, pow•r lk in transmission, and weight peualtims dut to li*k sins. At high Wnapaturcs. low
No part of the structure aUNOud ulks any pwmanet set or experience any damaW whea. subjecl to applicable lest loadts. 3. The structure sust not fail at the ultimate load. Wher applicable, ulLmate strentl of material sh~ould be detumined by bend••g or trusonal modulus of rpture. '6
k
AMCP 70152012
J
fluid viscosity can cause internal leakage and slippanc Compressibility of a in pmsacutr.advls fluid increaaes with pressure and temperature, and the resultant loss of voueoutput of pump6 is a compression offudprovides amass-spring condilion that can limit system response. Sut-cessful operanon of the hyrui ytmthroughout the design upon the interaction of lemperature range depends the hdalcfluid w~th all ohrcomponents of the sysein. sigt a" cnsdr h temperature distinThe d buinwithin tehicprinorder to achieve judcios lacmen o hyrauicsystem plumbing andompnens i th colerregions. However, it may be neressary to locate portions of the hydraulic system in high-temperature reions of the airframe. Radiation shielding and proper ducting [or air cooling may be nectssary for such equipment. The design of actuators locatod in extreme-temperature areas should be such as to provide for continuous exchange of fluid to assist in controlling temperalure. This may be accomplishe by controlled internal leakafe. which likely will be inherent in flight a lse. rim jxu-iali cnde control sacsiuors, but & be given to utility actuators, which may he subj;ected to extreme thermal environments for extended periods of time when in a static or nonoperating condition. Although intermittent-actuating systems must be desinedto perae wth oldhydruli flid, desfinetino opwerateauwithscoldsystauli w~idc their normal operating temperatures rapidly regardless of original ambient temperaturc. Hnat cxchanger and viscosity considerations favor vse of the Opg*
-
MLifiS5
tcm.peom.wa
thav -.-
m- -
s
circuit. However, this rise normally is dissipated as the fluid Passes through the system. Hydraulic system design begins with an evnluation of the system as it evolved in the preliminary design phase. The production configuration requirements and the ore cempared with the preliminary design. In addition,a ncsayhngsrcincorporated. ftesystem sUal be made in final exmnto order to ascertain not only that it meets the procuremerit specification, but that it isoptimum with regard to weight, cost, and performance. Ml L-H-5440 and its associated specifications will be p~rt of the procremnt~ specification. The subsequent discussion covers the major system deinaasicudgnmbrndtpofytm. prtn rsur retof~idmdaslcin filt~muon, fikiims, powcr levels and transmission system c;Aimization, and heat rejection requiremitus. "Ahe starting system isdiscussed, as are system analyais, including failure mode and effect analysis (FN'iEA), and reports required for meeting the hei141LWl~iPu~i~a U
The component sclection and design rcqui remepts also are discussed, as arc system installation design rciewt eea adteso oddsg application to components or installation. The hydraulic system is influenced by the size, and complexity of the helicopter and by the procurement specifications. Following is a discussion of the critical elmnsohyruisytidsin 0-2.6.1 Sanivablity. Rediabiii), sa Sakfyt Trade-
.ave
without encountering fluid breakdown awid excessive wear of movins pans. For MIL.-H-83282 fluiid. the normal upper limit it 275F1 maximum (Class 11 system). However. high-temperature design t~rds to impose a cost penalty duc to tW. attendant requiremint for special materials, along with a reduction in system life becaute of the reduced lubricating properties of the tlauid. In many applications, normal heat loui from lione arid components are adequate to "mantain hydraulic system temperatures within desin limits; thus, use of heai exchangers is uin0606Wy.than For analytical purposes, the assumption of uniform temperature throughout a hydraulk4 circuit Usually is quite accurate. When a pressure drop oc-
The problems of designing for combat survivability and for system reliability are similar. Redundancy of systems and components may be requirrd ini either case. For suivivability, the requiremeot may be a tolerance; of two hits anywhere in the flight control system without loss~ of the capability of returning safely to base. The elimination or reduction of the possibility of hydraulic system fires associated with incendiaries also may be a rcqt~irenent. Reliability aspect involve k~ses due to equipment failures. The loss requirement may be tated in terms of "no more n- kmas per 10 ).000 noncombat mission&. The component and/or subsystem reliability requi.emenss may be stated as mean time betwtsen failures (MTBF), which must be demonstrated by *jysii5,
temperature rise by 70F per 10010 psi drop for each
order to establish the basic system configurations
) curs without external work resulting - i.e., losses through orifices and tubing - the hydraulic fluid
test, and service. TrAde-off/optimization studies are required in 9-7
that will met the pretviously mentioned goals.The a]terat.aiva t iit may be considered arm 1. Two oi three independent. normally operative systems 2. Normally operative system plus emergency backup s~sians. the backup(s) normally being inoperative until loss of a primary system 3. Normally operative systems plus hydraulic circuit breakers (HCB) 4. Intersystem switching for redundancy with appropriate provisions against toss of fluid in the newly applied system if first systan failure is dee to lows of fluid S. Combinations of independent system plus intersystem switching and hydraulic circuit breakers 6. Use of armor in conjunction with system redundancy * 7. Reversiobi to manual control where applicable and possible. For example, computer analyses indicate that two systems with HCD incorporated are nearly a%survivable as three independent systems, and that significant weight and cost savings may result from this approach. A complete analysis of weight and cost imUh JV*UIUJ 41455 Y~..@W~ IUUS~S5~i5I7
replacement, filter changing, and flushing system maintenance. The RLS concept can sense and isolate a leak of any magnitude, thus reducing subsystism component and/or line leakage failures that normally cause complete system loss. RLS is detailed in par. 94.5.4, AMCP 706-201. 9A ZSse wck -. 1 SseSwcIgCocpps When normally operated or passive -backup systems are used as backups for flight control systems, intersystem switching is an important consideration. The switching function must be very reliable, and must provide for elimination of inter. sysem leakage as a steady-state or transient condilion. One method of accomplish~ng the switching func. tion is through the use of a lapped tpool anid sileeve, with pressure and retuirn of both systems on the samte spool. To minimize intersystaun litakage, only the returns arm associated directly so that diffcmrdtal prepsure between systemi is minimized (Fig. 9-12). A socoiid alternative is use of chock valves to sej"-rate the pressure side oi the systems. A powerW (p~ch~i: and spring) shuttle is used to switch the
L--ad1
d-icted as early in the program as possible. Close coordination is required with the procuring activity during this phase in order to insure that contract re.
quirmnents, are being met and that the design isapproved.
The combat safety aspects arc concerned primarily with fire resutsing from battle damage. Fireresistant fluids, are. developed and xhall he considered. HCB conci~pts also should be considered as to their lire minimization impact since they c~n limit significalitly the amount of fluid dumpedt into a fire. Noncomnbat safety aspect are focused primarily on trwn maintenance aspects of the system operation. This ~ttis covered in detail in par. 9-2.8. However, a ai-fl~ght cuntrcl' safety requirement dictates that if there is a complete power fatil~arc, at least one system shall be driven by the autorotating rot-or.
'ucomplished indepoendetntly. This approach may R04 be acceptabie under cetnain dynamic conditions of system operation (Fig. 9-13).
LMMDfrVkL#M @S.CUt.-., 'S'cuEMEN
Figure 9-12. Comblowl Spool S.Ik*ia Valve oz *ME
9-2A.1.1~ Reservoir IILel Semiftn
Reservoir level sensing (RLS) is a technique which uses the reservoir to optrate subsystem presure shkotoff valves mechanically. 'the reurn portioan of the system is isolated by a check valvc !mcated in the reurn Jim,. Soth ihe mechanicaly operated pressure sbutoll valve and th-, 6-turn check valve shel befu ioacaid as close to the central power source as possble in order to pro~vide *he assximum central powzr source protetlion. RLS allows a reduction in putrip
T FI~Wm 9-13. Prenw Cheek Vahns Fte feww., UaM sSwISaIg
third alternative is the use of check valves to sepatate the pressure side of the systems, in conjuncline. The relief valve represents an energy loss in the normal system, and acts to keep the backup system fully serviced as a part of the normal mode of operation (Fig -9-14). Upon reversion to the emergency (backup) mode, the relief valve setting ishigh enough so that maximum generated return pressures will not operate it. Where this concept is used as a part of a dual system actuator. anticavitation valves ire required in order to keep the relief valve setting at a reasonable level, A vsriation of Alternative 3 is the use of a mechanical locked-out relief valve to eliminate energy loss during the normal mode of operiation. This isaccomplished by using the last portion of the backup system reservoir stroke on fillinS to hold the relief valve off its seat mechanically. The initial sttoke of the reservoir allows the valve to reseat and preserve the system inteigriiy upon reversion to the backup system (Fig. 9-11'.
SYSTEM INLINE RELIEF
VALVE SYSTE
-"
CHECKVALVE -
H
RETUA#
~vent
R(SERVOIR PiisSok FRO
The pressure-sensing concept can be uscd in conjunction with switching functions as a means to preswitching a good backup system into a subsystem that has lost its pressure vessel integrity. A time delay of several seconds is integrated into the switch; ig fun~ction, i.e.. the initial motion blocks both
SYSTEM
llp,. 9-44. Pressure Chick Valves Plus aI&&e
R'tew.
RE
942.6.11.3 Rc~uru Pressure Siessiu Return pressu~re sensing (RPS) is most appropriate for use wich full-trail solenoid selector valves. The cylindes- port(s) is cominected to the retuarn when the valve is in the dc-eniergized position. A springopposed, pressure-operated shutoff valve is located in the system return at the valve return port or downstream. Normal system return pressures, by design. overpower the RPS piston and allow normal valve op'-ration. Combat damage, or component or line failure resulting in external leakage, causes reduced subsystem pressure because the return check valve prevents reverse flow. The spiing then operates the RPS valve. thus inhibiting operation of the solenoids and, in Ieffect, preventing use of the damaged subsystem and loss of the complete system.
NURM~AL
VU6SYSTEN
I
A variation of Alternative I involves use of poppet volves and a mechanically operated carn in lieu of the Pyrotechnically operated valves also may be used to provide switching. This approach generally is irreversible, and replacement of the valves is required after operaticn.
BACKUP BACKUP NORMAL PRSUERTR4 PRESSURE
NUKKAL RTR
ROWh Valve
MANUAL OPERATION
ft~dt
CAPABILITY S14OWN CAN BE MOTOR OPERATED
TO 11Cm R TLIR1h
ftEORIML r REMAVONSUBSYSTEM
SUBSYSTEM PRESSURE
RETURN
lipi 9-isS lasiL,. Mecimslcuy Ioce~d-eut POW..
Vshe.
Fip, 9-16. Cnm4*@mi$W Peppie S.whhirig Vat., 9-9
A
4.
AMCP X*202K 9-2.63 Selection of FuidW Medium the normal and the backup systems. During the time delay, the pressaure-sensing system tests the subsystem MIL-H-5606 fluid is the most commonly used and mechanically inhibits the switching function if medium and bas presenlt widespread usage throughthe subsystem does have dn external leak. Fig. 9-17 out the mailitary world. contains e schematic of this approach. A synthetic hydrocarbon &-fined by M IL-H-83l28 is being considered foi use in Army aircraft. Its o~idralossnow 9-2..2 pertimsPrere 9-U. Pessue Opratlg Cusldragoasprimary attraction is that it is significantly k fles1amOperating pressure is a function of helicopter size, mable than MIL-H-5M0. Operatiogal charactzrcomplexity, and performance. Smaller helicopters istics at ve.-y low tempecratures have not as yet been with manual reversion capability may use low fully established. pressures, such as 1000-1500 psi, without significant For a thorough discussion of hydraulic fluids, rafe& weight penalties. Generally, Jynamic seal life is better to AMCP 706-123with reduced pressures. Larger hehcoptcrs will require 3000-psi systems in order to atvain reasonable 9-2.A.4 Rletake~ ýCeeiamlswtIu) volumes and system weights. Syste.- fluid filtration and external contamination The impact upon development, qualification, test factors must be considered since they have a direct and maintenance equipment requirements may be a efiect upon the reliability jand serviceability of the hystrong motivation for maintaining 3000 psi or lower dratalic system. pressures. However, it is desirable to conduct a systmr pressure-weight-cost tradc-off study as a I-2.6.4. Flud FlihIIUIII mecans of determining nptimumn helicopter hydraulic Serafittonmhdsrevilb.Teeisymc cofigratin. ar.9-42. ACP 06-01. chide: central filtration, subsystem filtration, return contains a discussion of the pressure selection conitaon case drain crncepts, and suction lirane trtin uiderations as related to preliminary design ýuncCentral filtratio involves a pmwp~e filter, a return The basic selection decision usually is made inethe preliminary design phase, and a later re-evaluation mybe necessary az h'.4icoptier requireirciets ma te te hrdwre chage dsin s aiv~cesino phangc . th einawieinotehrwr
S,ýTEMeither L SYINRMAL i BACKP BACKP~SY~t H~tU~t S~E1~each _
vkt
ma ,rator toer *no a r ~smp cam drain fmm. Ybi method reqjuires that all critical coinpooests. suich a flight control actuators SW have inlet amcema. These screens sAdl be in the size map of W0-150 cron absolute. particle. in.e. purtides with two dimensions larger than the m~aximumi allowable Anil be blocked or prohibited from passing through. Subsystern filtration involves a pressure filter zhat is integrated into or is immediately upstream of component or subsysdtem. The retuea filter generally will be of a common sminle. configuration. ~Thecwsdrain from each purnp AaUhave aseparate
~~lThe
return case drain pliflosophyr applies to both of
ILCI6
SECTION
7
h foregoing categories. The case drain and rautrn system Filter functions may be combinsed withn a sigefiltet assembly. Thc assaribly may contain either a single- or two-stage element. It the elemnwrt is two-stage, the full or transient flow stage shell be per
ff
NOTE
VALVE
U()Wd
95u85Y,TEM
IF N*$LJ( wE!,:t
S
IEGIv
tIN
EXISTS.5mcosasltTecpbltofh
SE0SAi MILL ALLOW
COLVULON Or SasITC1,-
CU(C)IM~LCY
PRLEACImiRL SENSI PC SEC~iN
IW IfPZ!A
C(WEOtr IF~ U4.te
S.iTCsIsNG
Flgiuv 9-17. Svwkckdnm Valke
9-10
VE5SEL
HA L0 ;h'TGI'0HE1
SU'BSYSTEM
irn
IN TIAtLF85(
TEST Ai,)E
lt)ad1elo
lwo
bo
irn
oric,1
yaig
tesd
e3o ilk9
etement may be a compromise between the two. such as a 5-micron absolute capability. Suction line filtsrai~o Severally is niot utecd bo-Ause the suction presurc capsibilitics of pumpsi dtmand that the suct4n k'ne rressure drop be Lel~i to a minimuni. This means &suction liue tlentat would be qauiic large: andt. thee'ofore. buSL, and heavy. 71w Weight Penalty usually Is We5CCpflabic.
'
706-202,
_______
Filtration islkdher diicvcsad iý pi:-s. 9L4.2.2.3 and 94A4.5. AMCP 7a6-.205. 4-4.42(kud p4-n&t Fptnmioa Grouand cart gsr'cmaly trAxwp.nrte their owns fiIwicron absotratimt vysans.k whict mA-y tst of .%4-5 lute capability. The heclicopter fiyýAm s&11' moorporN.e provision for introduction of th.-ground cart fluid tbrough the syston pecusure filter(s) as shown in Fig. 9- IS. This provides for helicoper systaw protection if the ground cart filtration is defective. The two available altcrraui~es in filter elaenuts an the noockanable tLrcwiawayz cad ato rcusat-k. cleanabl types.
-4
94.,A.3 si~b L~din Tht required rihnttion lewd is defined in MJL-li54W0. which spacifie that tkn riatration moti at beast
6
rcjuiwiaentaa of MIL-F-12Z15.
4%th
The elmtefL, may bc eidwe noacleanabe :throwr-
vtays or clew2WkA, rv;2t~r "tcnertts.*Theehasiat anUZ tak c bic&ted frogi, mtAl vcn-u-tý
) ae
twardislcuirgoai; I.e.. tMk-t~i4soc relatvc dcp~h in tbz li'il TWeAU Ik-r lijgio4.11a thtowaaty Cknentu U*aItdly thei~i~oeniotn¶ ~s reLia~ifl pniwd* Siax.
-.
fli
if'
V1buile bosh nwny
4
the sa~ g&us b-adW wu cr"
12n2V-.
9-2A6.5 Flitkmp To facilitate installation, system components and lanes must have disconncct-connoct points consistent with specific kicoptcr itistaflation rcquiirem.ois. Searable freAable leuOLAirn~b 111e ret-UiW for cast riwmovall and maintenance. 1-owevcr, iij Ddafoy cases. permetnent fittings may bc used Iii joirijap runs of tubiaiV whc-v accixu through fthinstallation isnot rcqurt4. Mr. 9-4.3. AMCP 706-201, contains a dis-
cuWoo; a.' availablc fittings or both types.5 xxvewrveral reusable rfiting alttrnafives, iniuSa b MS ftlamA~is &jza AN Cawrst stazidird: that
MIL4-IbkW0 coven the reusable fMCitt~'iv
Icc-
to mom fsttings and aircraft. Howecver, the sýNciiic
of Lim ekAmue.
isaaar arc%~ in t6w helicopter must be analyzed. suit
944UA4 tJ&I#4* Ccw, EApcrieisuc LUa %lscwuttlia, %ut~a~ed, 'am V6. awry T~iViýR0, to poor hyklýsiutk Vj~A~ oýisQ aM.t hW4. Dav to the h it Auik taai ArntcOLVW~ w naujitrwn *uii eawre". twl fiari*g by Was4, c~.u&itixma~
necessary in ordeir to insure that the rftting cho~w. wilt pwr-form saiTrfscto-rily in senioc.
V
Lv ohtoft-4I
P--etv
~~~baw i.5 aa obvious and dcruirae mzansi of mini-
addstional teseins must be specified as conisidcrc,ý Tht Itosan. fittinjy recently has tic-tomec availabic as
aanpnlo-n~
m
)ti.AIlNKU'SLh15
ticajo---i in V4s- 9-19. lIz advantages over tile stndi AN~'J bo arA4 filt;-hý comnulml&ior tnc7idt a okn crw ln lri supcrk' sci tksi.n vi inaru cr
I
LOKIfiG FEATLIif
Vtn
L'
OW .F
t 4 ku*F
FLA
V
V.I (;M, V:
rWs 9l-M t& WyawM Symo isn (S.~ac -A L',-ewh~eV
-
1c.rt;1 include thc; Fv ..ioz.TsAlso "Dyn~ssmwlx said avuca' others and ajxz deirybtvS ii p 1.4.3. AMC!' I06201.
firm a erticlas that arematc tlWm the nominal r.%.0r~
CS16W fflf4y
~
another aletnevtive is tht use of rotary output$ in conjunction with c~isnk arms for conversion to linear motion in lieu of liner actuators. In any event, qualification and environmcnmtz Rest requiaemcns thal arm much more stringunt than those in present Military Specifications must be useý-A in order to asawestsia that new equipment will ptgform satisfracorily.
WijA.,s
r" 4411Po.
t
w, a
,'. 9-Ul
The use of titanium is being initiated in Iii zs and fittings with significant weight savings. Techniques for eliminating galling between fittings have been perfected. A weight-cost trade-off covering the use of titanium may be performed in tht; final design configuration commitment phase. 9-2.6.6 Dymatinc Fluid Coumectlous Swivels. hoses. and coiled or cursion tubing arc alternatives to be considered where relative motion between an actuator and a structure must be allowed. Because these types of fittings may impose weight and cost penalties other means of solving the relative motion problem must be considered seriously. The use of an articulating link between the structurally mounted and immobilized actuator and the~ moving coiairol surface or other subsystem is an alternative that has been used frequently. Fig. 9-20 isan example of such an installation. Par. 9-4.3, AN$CP 706-201, preaeis information coverinr,dynamic fluid connections. If hoses are the only alternative, close attention mus! be given to desiun of the in.tnliatin. The LOWminmun alowblý bndradius of the hose must not beexceeded during motion of the actuator. Thc hose X'3 postin rlaiv tost. ture an ra nthe actuator during motion must be analyzed in a layout daisioretodetennine thtinterference does notexit.In ddtio, aprteciv coermust be installed to protect the hose from abaindamagc due to vibration and g's. which can cause the hose to de.. fxtoutside its normal path during motion (Fig. 9ýrl
Swivels may be used in conjunction with hoses, or as scissor assemiblies. to rtsolve relative motion problems. The swivel assembly rigidity minimizes vibration and acceleration problems. However, the Joints involve the use of dynamic seals, which can result in nuisance or catastrophic leaks. Waing links is the preferred method. Peak Power Leveds The determination of ptak power levels, and, therefore, of hydraulic pump and Iiac siz~e. isdetailed in par. 9-4.6. AMCP 706-201. It is desirable to develop a precise knowledge of the parallel-series operation of 2he subsystems so that adequate performance is attained at minimum weight and cost. Undue conservatism in hydraulic system design will penalize helicopter performance, insofar as weight and power extraction are concerned. The same techniques of powcr system analysis that. are discussed in par. 9-4.6. AMCP 706-201. shall be used in establishing system peak power require9-2.6.7
man,% In
wrldifiinn 2 mioitinn nrnfl" on"lvtI"
co Autdto determine total energy requirements and the system heat load. Fig. 9.22 is a typical example of a mission profile requirement.
9ZU
AI s sieSatu It is customary to use hydraulics for starting APU's and turbine engines, since the hydraulic starting system is self-contained and provides a capability for multiple starts. No external ciiectrical or ground hydraulic carts are required, and a hand pump allows recharging of the accumulators for stubThe basic hydraulic starting system con-
SWLVLIKSsuccessful.
or am Argculathg Lld
anected
I(UCOPsists*ofTair-
1111
ýIhTA
9-12i
totothat
E CrSEQ
FIgu re 9-20.Use oamAtcovetbi L61
W1,Jhe.
Fiur
1ose
OU2Tyil
COMAT
CFTS
R Biei 1ACKa1'o E I
AMCP 706202 a manually operated selector valve. Self-displacing accumulators can be used to in.ure that the asm)ociated hydraulic system reservoir is kept to a minimum size and weight. The cold start is a primary design point due to high line-loss charactebistics with -65*F fluid temperatures. The motor requirements for equivalent warm oil output torques are not increased appreciably by -65"F fluids; however, flow rate will be greatly reduced until fluid temperature has increased. Instantaneous or fast opening of the manual control va;i'c can cause hydraulic motor shaft shearing or oth-.r damage duc to the high-p asure shock wave. The control valve design should include features to provide for slow buildup of pressure (0.5 to 1.0 sec is reasonable). At high temperatures, the extra fluid enrgy available as a result of decreased line losses may require control if the APU or turbine engine is aoceleration-limit,_ due to a characteristic of a cornporent or the basic gear train.
9.•
\
9-2&. -
System Heat Rejectwo Claracteriscs
,,,-,o•vci
t--:y,_..
.
-
,,-nr MA
--.
or less, and low flow of S gpm or less) generally do not require any special cooling equipment. Large. high-power-loevl systems operating in relatively warm ambient temperature regions may require heat exchangers iii order to maintain fluid temperatures below the Type !1 upper limits of 2750F. The performance specification for the heat ex. changer includes the following requirements: I. Media maximum inlet and outlet temperatures allowed 2. Media minimum and maximum mass-flow rates .nriMaum IIJIMU 1WrMAIMlaUnM mnCLi and outict allowable temperatures 4. Hydraulic fluid minimum and maximum massflow rates 5. Allowable pressure drops in both the media and hydraulic sections. The heat exchan•er qualification test s"all specify the following basic requirements in addition to a demonstration of performance: 1. Realistic impulse testing as a pressure vessel so as to insure adequate fatigue life 2. Environmental tests, iacluding vibration, shock, and corrosion testing. 9-2.6.10 System Awslysis "The system performance analysis necessitates definition of the requirements of each of the subsystems in terms of output requirements and resuling "input needs. Flight control subsystems and utility subsystems are the basic divisions. Par. 9-4.6, AMCP
706-201, discusses the analysis or determination of power requirements for both types of subsystems. Where there are several subsystems, the maximum simultaneous need must be detc; ntined in order to define the peak output capability. As a part of the performance analysis, line-sizing criteria based upon maximum allowable transients must be used to insure that MIL-H-5440 and practical limits are met. In order to insure that performance goals - including combat-damage tolerance and single- and dual-failure tolerances - are met, a failure mode and effect analysis (FM EA) must be conducted. This sAW. be conducted in as much detail as is practicable. This analysis sAWll be prepared in accordance with the system safety program plan of Chapter 3. AMCP 706-203. MIL-H-54,0 defines the hydraulic system dat and reports that are rcquired. including: I. System requirement studits data 2. Design selection data 3. Developmental data 4. Production data
5.
Schematic diaRram
6. Hydraulic system design report 7. Hydraulic system nonstandard componenrt cross-sectional assembly drawings. The requirements of the helicopter detail specification, together with the MIL-H-54,0 requirements, shall constitute th- basic hydraulic system report requireten,. 9-2.7 HYDRAULIC COMPONENT DESIGN AND SELECTION The detail n for the various components needed to complete the hydraulic system are discussed in the paragraphs that follow. 9-2.7.1 Actulors Actuators may be separated into two basic classes: flight control and utility. These two classes may be divided into rotary and linev- output types. The flight control actuaors may be dual- or single-system types with manual, mechanical, pilot control input signals. In addition, some units include an integrated control augmentation system (CAS), a stability augmentation system (SAS), and/or an autopilot system. The controls - manual and electronic
-
generally are
combined oith the electrical control mode, either in series or in parallel with the manual mode (Fig. 9-23). The series mode provides limited authority and flight control surface motion without motion feedback to the pilot's control stick. The parallel mode or operation provides unlimited authority: in effect. 9-13
A
ELECTRICAL CONTROL INPUT
I I
INPUT
(A)SERIES MODE OF OPERATION
.4
rrenDAr
I
I LLULIflUII LIIIRVUI..
P1 INPUT
SEPARATE PARALLEL SERVO
ACTUATOR
(B)PARALLEL MOOE OF OPERATION Figure -23. Eaaks
f Puralel and Seek. Couhol Med..
dragging thc stick along with the surface during motion. The augmentation and autopilot modes of operation use clectrohydraulic valves, which receive electrical signals generally in pvoportion to the
load requires cons.de.ation in the decision on type ol valvc to be employed. The master control valve that accepts the iwinual. clectrical. or combrcd signal is a rour-way. closed.
magnitude of the desired surface position change and then transsate the signal into ahydraulic command to
cenler valve (Fig 926). For balanced-area cylinders in the neutral position, the cylinder port prewar.. are generally established at one-half the system pressure
the actuator. The electrical fcedback signal cancels out the electrical command signal when thc Loritrol surface reaches the commanded position. Flapper
and nozzle or jet pipe electrohydraulic control valves are used. Figs 9-24 and 9-25 present schematics of the valves. The jet pipe is preferred because it hai an inherently greater tolerance to contamination: however, this type valve has a higher leakage rate than the flapper and nozzle type which increascs the hydraulic system heat load. This contribution to the heat
at the no-load positiwi. In unbalanced actuators, the cylinder port neutral pressures are unbalanced in proportion to the area unbalance as requirid to obtain the necessary force balance. The spool of this servo is subject to jamming, particularly by thread-shaped particl's which may have been scraped from wrfaces
of co.aponents by abrasives (predominately sand) which arc introduced from she enironracnt. Sesise scrupulous care during maintenance and servicing o
K-
AMCP 706-202 preclude entry of contaminants. the obvious pre-
ftmom
MWL O L
IIALI]
ILJTO
use of depth type filters which trap these threadparticles far better than screen type filters can
PAshaped
JET
Where the electrical and manual signals are mixed,
linkage mechanism is used to combine the signals at
7a
the master control valvi. The manual signal comes directly to the valve through an appropriate linkage. RECTURNI CYLOC1LNThe clactrohydraulic valve output flow is directed to a small auxiliary r~m that, through an appropriate AT'Lucelinkage, can move tht master control valve as ~rnrMciuodirected by the electrical signal. The utility cylinders generally are relatively simple
SWV9 C,
Fpe9-24. Sebemate of Jo Pip E~r*
r
and do not include the control function of the sub9-2.7.1. Rip-stop rrotectiee When dual actuators are employed, or two independent systems are used simultaneously in a flight
c Kg
s
H
UTOcotl
Ill
fl
II
~ ~
____
***~ ------
kuyl
oto cutr xmineidctsta aiutic cutrepeeceidctstaa failure may origin~ate in one system and propagate
into the other, resulting in loss of both. If control cannot revert to manual mo*i in such a situation, the helicopter can be lost. This potential problem can be elmntdby use of the rip-stop tehiuwhich requires the use of separate pressure vessels. The vessls can be joined by fasteners. brazing, or other ossiblityof a fatigue failure causing both systems to
9F1
b lost Theprescrnt state of the art allows use of this
technique without significant weight penalty. There-
75fore, it shall be used where possible.
PETUN
LAMA
FlIe 925 Sckemtkc of Flappe Elecetro-hydraulk con"a Valve
P1E.SSURrRETURN CONTROL SPOOL
PASSAG-ES TO CVI NDER
Flgmwe 9-2C. Typicall Master CoatrfA Valve
9-2.7.1.2 Endurance Testing Requirements Adequate endurance tests are essential if the Fkctuator ZAr1-4" 1;fictimic is to bc satisfactory. Tw.-o important design considerations are the dynamic seals and the integrity of the actuator as a pressure vessel. The endurance test program shall provide foi a rigorous test of the basic and detail design concepts. Toward this end, the usage must be defined in depth, where test requirements ii, addition to those s'-ecified in the applicable Military Specification are desirable, they shal be specified. The additional test may include environmental, operational cycle, and/or prcssure yessel impulse testing. 9-2.7.1.3 Seal Alternatives Military Standard (MIL-G-5514. etc.) 0 rings, backup rings, and glanJs arc used in dynamic seal applications. The backup ring is required for pressures above 1000-1500 psi, depending upon the specific application. Frequently, seal life requirements are such that special, nonstandard seals are 9415
-
/
ktA'IV
aiu
vtso
ow,~r
sasacds
4..,AP~.CFWbM cuIdi p~r 944.6.This CUS.d
Pverfiv4 Y '1 '.This
A Lixful tcchmicrut £ involcý the ti.- of dual seals with the sw~Qivfl hetw-41.lm !sc als vente-d lo the return side: of ithe iystcm. Thc first si,;Ze ii, then the highpressur,ý -x) uitd thie rtcond stugec must 3cai only thc li cwr pc~~lrvausue Revricior., F~a he used in the secion in order tc knbl-it large, ý.hort-circuit laa flows, UsX 'IoY t -ruly oundtil sCHIS iNd OMw tehniqUe. lnvih'e'. us of tIo scars in -crics without venting, irt betwon. The &ksi1'ncr shoujld, h, wever. insure thLIt fithid cainr-ct be trapped betw-en seals or thAt !srnleffects will not cause fa~nure. ring seals may tyo citiner fu tndn Backups~~ stanardscafedor nscirfe (slid. Te uscafcd btaciupr scarfb ued or nwvard psolid)lTe lieu un f d baciup hal whavvr beuse pssiic n leu f t,,e standard scarfccd types. Srirfacc fin.'sh ii important to seal life, and should bt; kept bela.. 8 gin. A )&biiir. finish may be accepL talcl G'r shor-We ap~atos g.eahgrerally in use are the MS standard o Static rivL.s with N'6 standard scarfed backups. As with t-ie dynamic seals, unscarfcd (solid) backlips are preferred %heic the diameter is large enough so that they can be installed without damnage. A It)- to 32-14in. Finish is requirea f- aci O-ring gli.Ai2g 'Thetyps c~rrra~ oseasir us ar dicused ,-i par. 9-4 lv.6, AMCP 70i6-20 9.2.7.1.4
Natuital, 2nd Stress Conjideravio-is Actumv: stx._.cos.-
-. a.am.
ni;;
coi-ar.
(AV~J
stres us ri~e & an pprprite :'ienils ith provnsesscorosin rsiitnceandpreictble adeue finternalsand kept!inl &!iabc raiu is~high~ radis i hihlyderrabe. hc inih m.~s hrkep in the range. of 12 rms or belcw. A significantly rougher fin~h geeros. ~nofi~~ adi by ntrducng any no-rdiu s .es~ra~srs.wa; Sto-ls gcncially hive bctite, more cv:isistc-it falijzue-litec charwetr~stics than duc.; a~urnlnur'l, ant_, -'relorc. sho.-i be used where- long life it mandatomy arid fliglit sF4!05 importact. W.ZL-C-55O3 is applicab nth~s zem:, ane cointaia-s detail requ;rcinens in. leu'~ n ie~ual 7 ht !apoed spoA-.Aeeve zonitrol valve matutiats may Lic cither SAE. 521(V) series steel. or corrosionrcsisy.snt ste-1i such as T)pc 44'1C stainlcss. Tinesc rn O have p-odcn swifacvury in scrvicc usage, leat 04.. C slainless must x -.old stabilized to ir.ure iaime'nsilor.' stabitty in cr'kr to a¶'c~t suosuq.uunt ,ammi~tt.failed
-..-
em
eurwa
discusuon applies primnarily to flight control actators with integrated CAS. SAS, or autopillot medes of operation. For satisfactory service usag, .it is desirable that the nccassary control linkage bL imniersed in fluid inside the actuator. Jamming. corro'. sion. high friction, and unaC4Cepabk COntrQI systeM slop dmc to wear Art problems that arc eliminated and/or minimized by this niethod, It also allows the use of a rotary low-presaure dynamic "ea as the
access fromn the cxtcrnal manurl control system to the actuator internal linkage. Such a seal is considered superior to a linear seal Wcaust more efficient sulainL minimizes introduction of cotaminazlt to tbtsel a fiamdgsrou Mechmnicaloeri appropriate margin or yceld/faillurc stirength in the control linkage. Reliability of ovearid swoss depcnds upon the iieverity 4~ the jam, which, in tutn. varies in some mtanner with size, zhelpe. and material of the jamming particle. Dependiiaoe on mechanical override (as the sole back-up for r jammed servo) is not an acceptable techriqic ecrp: for stmall (1,... contrai fotce) helicopters. Hydiraulic rWundancy is the presar~ed mvethod for lu~ger helicoptera. The linkf-ge and the 'ercc: input ;;apaoiliky hall be designed so thL: cmii)p that may 3et to the control valvc oin be sheared, thu~s a,.oiding catastrophni jamof the a(.tuafor control. ricnerally. a I000-lb for'~e at the valve spool centerline is used to define the .zltimai, loa:1-carrying capability of thec control system linkage between the pilot input point and the valve spool, Thi; is conserviltive, butl usuaily dcfleclion and other crite~ia are such that no weight penalty is iavolyv.
For stable COnt7ol system operation, both with power on and with power off', actuator stiffness can be important. The fluid spring rate capakbiity of the actuator is a prim,- fector. because it is generally the sfetpncno h oa can pigrt aa bilit). Adeqtatc stiffness may be attained in several s. Thc. aottator artr. may be increased beyond that required f'or aerodynarmic pi~rposes so a! to increase thc fluid spring rate. Another a~ternatk**'c is to use hydraulic damp3-r: in conjunction with actuato,copability. Power-off damping may require the usc of separate dampers in an; event. Dniuretl actuator operati'on on a single system may require- hydraLlic bypass or anticavization featirre. This prevvits thet pimping of fbuid back to the s)stcr reservoir. a.-d a resulting requircricnt for a lar 6 zr, heavier rescivv ir. Aanotlaer benefit is a sig-kificant redurtion in thi. energy rcquirAd fioin the rcmairning operativi: syistem in 'arder to virctilate. fluid in the syatcm. Ariticavitation can be acecrnplisho-d
eih by. a pmtHps bypm valve or by de vahvs W4 &W rlturn So~w to Sbr circuit %becmW~ valwl. Fig. 9-27 ,oniam admmtatics for S to a a . Ila• cwý an isdisiM " valve €Omltn-mor is ¢mar •tisW to 4fi tsa fluid in reained within the 7U re sdltir mus~t be hWSWc than We•
•'r Sicltua~ir.
6e
flow prmu
thatt ca be developz during 0054do. Th11e ant of auui g sW be consulovd wlozmr
Faedboc for positivi control mry be altziw~d in• lt kw thzee tvays: with a fixed inrrdi. %:ith a m~ovable twrel, oi- ekctrkally. Th hm•', berivi in.volves s movi rod and an sciated mechanica hnkw• interconnect tdwen,the movinlg rod *ad 1k¢ pilot input liskaage "Thisapproach Aiiow$ thec Uft of" direct tubing cim.mcttonz sincc i€.ir can ýc little or
Ike I•;aura-.lo
no relative motiors. The r/noving barM• giv•s direck ;*~ba", but require ho , swivel&. ot oftr ur of handling• the motin. The third atlminativl: in-
W
WC~v" tM uu of tramdamn ill onjut•'tior Wilh cklctbical eiccontrolF . 9-f2n containl chemtk.bs attai al-
anstf irt Wd.
ka can exist. Snubbing reduces the m cause op a presvlre inorfica-
Lhavalves t
taken into amount dshor
cnaciu desgn
a
spptowaha.
ahd led. vaelo Fig a ht/dealaa cffhmtivcex=s Uoe of an inl, check vatvl iThesirable, particuade-off is apuipr to daeso the optimum aparly when ainr dual actuaterd or ame used, If the aerodyShe maiUmE REow• DodaIo FERENIA[.bck.btreqirspesswvd other mears. aa rcash. ivmm load durintr single-system overation can ot. poweriptum, the tn e ua of ppoa a k valve will allow RETURN holding cfthe position involved when the first sysorm fl Ir tfailure occurs until the moment or icad can be OW PESSUEDIFERETIAL PRESURE Te u RELIEF VALVE laa --. IgeT -.l c t ghvante tco that m--s- be tan int i
L
and tet. ro
VALVE
iTservU , -ottor a
un turret dcivc applications. Vanes
may be used ad combination actuaior-dampenr.
VALVE
ACT TO UA
fLA
reduce to a controllable kv¢.!by reduction in heliof saedli r th motns. or motor type of "otary actuator may be iopgars. u advanaan in hcdg A obviousaapr is
JR•
R
PILOT
MO I E
VLo
(A)) UEOOAFCHECK VALVES
*
~
L~.(A)F 'RSU~
r
) SEOFCHCKVALVE -
S
IXABED BAFRPEL LO PRESSURE-DIFFRETAL•
E OEAE
M RESURR DIFFERENTIAL PREASSURELLOWINPUSSC=RE 0_
CONTRNtR
IIO
ILOT INPUC !C
TF'ANU0'EL G
.
()MVABLE BREL
*
)*
(B)USE OF PjflS.UP:- ,PERATED BYPASS VALVE Flume 9-27. i..srltketlppoeochus
,,.
(C)ELET!CAI. FEEDer€P; FIgure 9-28. Feeawck TchLa4ues
•,-l7
"
9-2'.7 Hydtauicw Pwps The hydraulic pump is the heart of the h'ydraulic dyasei. and. thereiore, i., of primo importance. It is, however, the primary soucce of ty-tern heat emsrgy due to its inherent inefficicnacim. The basic 0ta.-tatives for pumps are the rixed-displacement and the variable-displatement, constantpressure 47ypes. The fixed-displacement unit can he used with "opci, center" utility systems, or with relief vaives ot combination unloa&er valvc-awi~mulator sysiems, for pressure control. The fixvd-displaccment system is relatively heavy. In addition, 't tends to have higher heat generation levels than the variabit-displacemcnit type. Consequently, the fixeti-displacement pump is not in genecral use except ii lowpressure-level systems. The cinventional variabic-displacement pump has constant-ilo% capability through speedls up to the pressure control cut-in point. which is approxi-. matcly 2850 psi for a 3000-psi system. As the demand decicascs below the full displactment capability, the control de-strokes the pump until, at zero output fl~,
..
~.
--
iw~,,;
,.
*~~.
*
;n.
9 ...** .......
The compe ,sated pessure setting can vary from 2950 t3 3050 psi between pumps. Fig. 9-29 shows ihcse characteristics. Modificat.ons of the basic control art available, One variation used for horsepower limiting is the soft-ci'toff, relatively constant horsepower approach. Fig. 9-30 shows the control reducing flow at 1500 psi
soresult 7.
wad continuins on a linear basis until it vtadsx5 ttv' staitdaid sharp pump level. Thert, it. mm"n k, duwap cutoff control. S'aperiswczad an 5F4. 9-3f is a cons¶*flt-horsepovrc7 line that is tanrt tw .1v Pot'i cutwff and that, wken compared to the peal- '.onepowee of a sharp cutoff pump, shows a sqi~i~cw reduction in the power required to drive the 1%mip. This is a definite advantage whome the driving StPOWCI1 is limiltd, JIM Significant wVi~t "an ke talk' in, for example, an electric moter. 3loweý er, zome o' the weight savings is offset by the large li-a's reiuircd to maintain subsystem rates with 1530 psi avaitloW for load and line loss versus the conventional 315O. psi capability. Another type of c')ntrol ofiers a high rupt-nx capability because of the large control valve u,%4 with pump hanger feedback as required for~stwbilization of the pump contro; system,. Thv capabJity to, respond to both on and off demands aliowa the 4,Iction of the accumulator, resulting iin savings in weight, maintenance, and servicing. In addition, 'kc pump has better overshoot-undershout chars&;-tb,jitfir than has the ConkCntional nuimi-nL~jourL system. Tnereforr, pressure pcaks are lower ono' system fatigue life is imjpic-d. Pump case drain characteristics are iiafltrti' teo satisfactory system and pump operation. The L shall have a minimuin case drain flow as a functior. c'a back pressure. This is required in order to climi..'.-.. pump overload and failure during operation in IN' compensated or giear-ccmpensated mode. An uppe limit on case pressure buildur. due to flow restriction is necessary in order to avoid L.s cr the system as a of failure of the pump or case drain a~;rr sa
.1i1111characteristics
~ 0
Woo
0flw
Mw
arý sI'ow'n in Fig. 9-31.
4000
29W0 ,W
"ov0.0 sFigur r 9-9
CASEIWAM L" %F C"W.CTLM5NC5
Hydraulic Pump Flow ib Preitsurv
Ciaracteristlcs
I' I
iigima 9-30. Hydraulkc ?suap Soft Cutoff
*0491 WA PUR0
~CtIAMCTWITIC
Figvwe 9.3). fydraiollc Pump Casm DFial FUs
Vamp ims&mauI"ti ame impxaf POW yM*Gieis adW pump W The r oo r'
j
t t pros-
*wihu ia per. 9.4.5. AM2CP M-.201. and a hydrink pabolsies suppramo is shoen in F5 . 9-33.
.AW ha swb do fhe ~pw mag= dme m et wms wsk4 MiAW to lvi W rb free cmensated (nobew) subk& Fit 9-n2 ibm.. *uacsnhmkti systm deb. rqinisnama for nesavoir pmwmurs for tpsAI e'-wRiMg Itnepanms and mctisolne km ksV~k *&. £ nades suctisn kas tamer and
(26+ pmi) and loe imudane (la opening to ayesk) am required un order to anoin the masesary hjb frequeny repoese cbuareuurcics The bernic applicable pimp spairkacaioc is MEL-F'492, covtings varialadsiwry Pump requirmstws. MIL-P-7WI cover tuad~paee wnit
Pass Npakdwse cba,%nninics ma dthir cempusibility %A he sy~am dyiamic chnoracksars are aurfily itapthts RIA uisfa.cior systms and helP csWr pct-Vmaace Pimp syss-compr-Ability Lests smus be oreeducsed -4suily as possible in.the hardonwips cy&k ýad ban &Wweb li~ use dme to dos peep wftbowt' taua cm bast she pump p~hsies d6Zcbsrmatuic - -n exciiag It.9um*c- eaR duivihe aw PPmadnM Mae at its remooazun fu-rsuc. The rsudt i puhleakcs41:J4sr of maewsflfW M V&" tue,Vw-4j pub'twu Ileve
9.1.73 AcAmfar Accesulatoes me used for enery ztoage and/or for pressure Mtasmnt and pulsation attenuatbon. They may be of she pmston. bladdr. or diaphragm-in ps-fluid-s.earaion type. A vaniaton or the piston type is the sdlf-dicplacing variety, which avoidc any adverse Impact upon the ,esrvoir as a result of size or transent highvelocty motion. Seal, bladder, and diaphragmi leakage, and its cor.troE,rj te basic acci umsacor problems. The piston
44-mas cm,~ tn: 44 fr raeda!ýktJ vlbrzetivav- that ttcst %.No fVcb46lk ~ to situl fatigue fk N&Z' Ms k-eitminutn. fta . ý=M.-nt,.ukalaoi l-rv2, is un~a.Wetablc, there amt ivt. i4."in&Wciuive Lrxe,. Line lereibs may be vnodictr in thc rnostwinj arta sc As to move the rcsonataneýP speed point outr1 oZhc 6nr03 spvc range of shT pumnp since resonance can occurf betwtecr. tr; pulrnp and an inrne suppmnJoi. Them earcnvrAl types of suppressors. the Helmouhzl rmnalor~ r 6
-*
-rn
"~
_4
OD AD
0 42 __
3
10¶2
14
provides redundancy when the system is n~ot operating. Another alternative for minimizing leakage of gas into the hydraulic syst.m is use of redundant piston seals, with a ve~nt between them giving access to the atnlos,-.hele ItLia may require a tail rod and an extra srux. 'I fit tvolution of high-response pumps, and the resmnt ekir.,ination of accumulators, is a desirable wvay of rolving accumulator pr;oblems. The only basic #acimbtcnr-cqtiinig accumulators will be th,, engine,osinulator sit~ing hand prtrnharge arc a funtio of th:ý &rnperatwe rerage thirough which the accumu!or ttiUa be used. Th e fluqid volume available at tht; minn.:dut~i Fre~ssurt and the minimum operating temperu~ltt obviously mum meet the subsystem energy requiremenwts. Accumulators for tactical helicopters 11 he designed to reain their integrity when texp,.'scd ~.o gunfire. M'L-A-T498 and MiL-A-S897 cover the detail accun'ua'vor requirements, and per. 94.4.3, AMCP 706- 6.' , contains a di!,':ussicn of accumulator tech-
1¶6
SUCTION LINE IDJGTN, ft
Figur 9-3. Sadie LA ime Ls t Prusre Chwaraerhaks
-SWIA
N
Fkgur *-33. Hydrolic
rd~a.
SpaeSM, 19
.ip-
coiwm-arn ThMu of IE M chfiqlpy wh as KR " an eke RAmcvor "• iome prmp n"i in pwrentaw ramuI, re i OsI(rsf to p 9-2.6.1 oto & K LS "it mw be in,the 2.1-). Mnehhuicxt•y l vitumus "w e a" frM s• yma•ysdi nsdM " j pr-vide for thsL. - aswm. &adwws o MW grated into the &msp in the opfmo asomaan-:
9I&A
pTI
fovi racceptale lul
S dunp
leakae lt dynas•t
aye opnratioc. Tha- am everal typa
of r van•a th •i a di bootsa., tsparWS-mr frOmnsiscd, On OWxtasi-s e nd-ui puWSiaud. Time in aote dews in pair. ".41. type are diuSCu ts AMCP 7•6-20t. The varinch dnip reqalitrcs aW d tfutara of the resrvors ane pcumsed in MiM1-552C and MILy-R.W3•l. Lm ialy is defind by re levielmau The reservoir tie pump acio anrutiremuesa. Thee requirmaw
vd te Weson bth invove luidam,4erakm
tes
Z
•
'r
,ht
"c"lwtonumi-opavitation as the maumurca
a.
minim
e-.rt.g.
.o
must be harmacterisics points.fluThe panur involv bol~~ jsh pump y overranide thepsteML.-5t dea.d before the resrvoir ad pump suction be anWn and destpnd. MIL-H-440 deilin
~can
ge~ruisneionProvisions
seofc vere them. hove l rthes meotqiw n pumrp reno my ovrrithise Therseas w• in the ldht r m• n disaptragze zue-4'L reM not considyed dynamic inrthe sense of a pideon
ol
ailsMK
dyiwi t
pe'W
bile operatin,
tc. Funcfonuly.
oItgwwlht, Mi-
lt!.Me mug Ne
aid". d. dinS uue the specifilt suabs. 1on tin acdMnsticr. to Mksn far the smirch and hittn.pieoatsid PNO"MSM,
The sucion on•tlet should lie a. the bottom of. or - the cervucr, relative to rttha the miWdl of, no lhi nma full37veli. In the cae of conspurze rowt g-tr sacnt operation iV# Vals, provisMum bo n fo it
4V
nW
L.
t
flM leve? itdkr_,tioui sWd be an accordance with
ltyies r renn-v.rs thet high-pressure side she tranuienough to nusatain large passages stllW e on Yai.n n gnu dut to3750 rewtvoir gni. high-volwu chng rate from eiterading For bocw
shol be incolded for an air bind at the
hihest point of the ruearvoir as it is iTtall in the op.r. Manu..l-opratin capbility is required. COrw•dei
r tiSon
v4
egivr
e to ilred
povsiility ofo
asotnati bleed on startup in addition to the manual
seal. The diaphragm and bladdcr do have relative motion between themselves and the mecrvoir. T.he bootstrap reservoir does incorporate piston Mats. VWhea the weight pti.alty is acceptable, bellows are preferred to dynamic =sit for the low-pressure bide.
0-.j7.5 Preuaore R41f SUU c!ifvle rcrqie o rvd o Posjible pump con!'rol fnlure. The requirements are
level shell be estimated coaservatively and is to be
(M)s"efqulfe
Where dynamic seals crc usetd, friction and life expectancy am. primary considerations. The friction und in designing the rscrvoir to meet the accelera-
lion an.s anticavitation requii-ments.
The reservoir shall include, as an integrated or sepsrate demtnt, a relief valve for protcction against overall, in conjunction with acccpting the maximum flow back to the rcscrvair as a resul, cf syt'tem differential during operation. Fill provisions shall be in 2cc.rdince with MIL.-R5520. Filling is a function primarily of the type of reservoir. The opcr or air-pressurized reservoir may h• filln,1by hand with fluid poured froui a can. However, this method I-as obvious disadventages in that certaminated flud easily could be poured into the system. Separate xar-fluid rcesrvoirs rad bootstrap rescrvoirs cannot bc filled convenier-tly and easily b) pmt -ing. Thcy reqvire fill carts, or equipment such as ijand pumps that c•n generic positive flow at low pressurc This typ'e of filling normally it accOmnishvd through the system return filter. Thice is desirable because contaminatd fluid will be filtered pz~or to reaching llC pump inlet,
bined.
covered in MIL-V.8813, and a IWtliiary Standard (MS) series of qualifed valve is available. However,
avs
vial.Hwvr Is
MS valves are in separate housings; where integrated modules and service centers are chosen, the runetinc..!
.F.ns
okI
.:l ......
ntr.A .i
;into
the package. Consideration should be given to routing 'the relief flow through the hydraulic system cooi.r. This could prevent systc.m overheat under conditions where the relief valve an the pump compensator malfunctions. The staneard system relief valve rarely operates, and therefore can be quite simpie in design. Where high response, frequent operation, and/or narrow reaat-full.flow requirements dictaw a nonstandard design, consideration must be given to the dynsamic performance characteristics. This will include damping provisions, such as a hydraulic dashpoi. 9-2.7.6 Piewswre Regulatlon Presbure regulators or reducers generally are used to step the systen; pressure down for use in a specific subsystem. The detail requirements for these units aor covered by MIL-V-8tt. The ret;,cers may in4ckd
_--
rmitr valves for prosolicm of the sabsyum in the
)
Spressu'e
to limit Ow differmntal
eve tin uudMer fails. The perfaonanie or r*els.
presoure arogs the rilr elment o that system per-
tion may be absolute (based o atmospheri pmmuw)
formasc dqradatiou is conro•ed during tranient
or difnotWl (maaienan of a specific pressure leve above the return preusure at the reduce in-
Icw-Impwume opauton. The intmerctwons o; se reie
utalllaton point),
lickout, ditftential presun ladiclor.,an element dean-drty diffaemrtial prmw charad stica are
9-V.7
importnW. The varioum requircun
eRi
must omple-
meat each othar if theis Musi. is to rfmction effoctively. Fig. 9-35 shows a typkWl composite performmc curve. Thi cuv miust be ed as a tool during dWen in order to dsriee the varius performa mpkeand -enem they *IV n requirmu mentary. The reziui•r•eca to an emiet differential pr:surs indcAtor is coverid in MIL-F4813. and calls for 70* i. psi difi• ntial capability foc a 3000pMi syrner, element. ?Aimration of t fluaid rake me" is not acceptabe. This face mwut be considre during election
meet.
or the mdia to be ued. If neocssary, ad.*tional
MIL.F4815 requires that an automatic shutoff be I nrnviSM wn that *, minimrnun amnunt nf fluid 3 |s
tf•uing Ma 1nmAiain i&.
Thermal lockouts are used to keep the dif•erential pr•sure indicators from operating below a spwificd fluid temperature level. This allows cold 3tarts, in which the high-viscosity fluid would cause highpresure drops across the elernent, thus causing the diflcrential pressure indicazwr to operate. A time delay is used to keep the differential indicator from operating as a result of transient peak flow pressures. The contaminationholding capability then is referenecd to steady-state
9-2.7A Clhick Vales Check valve rmq•irancnts arc covered by NIL-V25675, and AN check valves are available thba nect Type 1 (160*F max) syttem requiremen•u. MiristurizcJ check valves ar• availab for Type IJ syzytc (275E- max), and provide a wcight and cowt nhvins. Nonstandard ca-ttid$c check valves that can be used in integrated packsgcs also are avpilablc.
~m.~
be conducted so as to coafirtm that the zdaux ".
;...-.tcLJs...
140...... .O.7
fl....
.--
A
*4.L
.#.1. Uhu
.a
*n"
without penalty to the system.
if
/
C
IP a.
H
or
G
0
.-
I
__•"1•1
60
L
INDICATOR
-
PUE40. FNG:
OPERATING
POINTIS D USI7GUN
40-MAXIMUM RATED FLOW,
20
-,
NOMINAL
-)
_ __.._• ....
• •0
F.•Z.sc(
CS CP4ARACTEMtIS
100
L
.140"
t
I
lOSEATt
.
v.1w. thermal
Adequate filtration is esmaial to mtisfactory operatic. of the hydraulic aystm. MIL-F-1I5 coven the detail requirements for the hnussap and claents. Filter eliments ar dis.cused in per. 92.6.4.3. The Cirt-holding venrs differential-pressure char. acteristc ar shown in Fig. 9-34. They arc quite nonlimnr. i.e.. a significant amount or contamination wiJl be uaiuw• before there is a notic•able inctnase in differential pressum. Differential prenure indicators af required in order to signal that the clement is loaded with contfmination and requires replact-
when the element is removed.
W
A bypaw rdief is ud
..
~ ~4r
CLEAN ECTENITICS
H
zak-20l'T t
7i) 80 9O100 110 120 132 cc 00 0 DIR-HOLDING 0 ll5060 CAPABILITY, ((11SO rbl
FuRpm 9-U4. Filier Elemmo Dnr-ho34wg ('her9a3.
THEEIPAL LOCKOUI
0
'EN'NC
4
IL
0
t:..LA L 5IIJ..Zj. ±. Z
S
F~lure 9-35. Filt*.r Ebeemi~ Perfer~mae TEMPEPATURE. T
lke Elsllnaimie
-ANGE
92
The rniniLturiwd check valve, with its inhcnnt bMw-innhtia. higz-rapwu chacaicriskia. caun be u-A] quait Liectively to kcsp trauskals oV one subsystem fromt im.'oeing unrnecessry, ptoentially diumaigin Maasses on Owc cib sub.1ys~erns. Par. 9. ~3.2.6.1 contauts further dascusauon of chock vaivs.
'Sq.
-4
A11-2.7.9 't~c
1
>4:feature
Prummwe Swhdss Tc pr-axts switch gencraill is used to opevnehL
sysian failsare warning hight in the'cjk TNe switches may be cn~itfivc to either absolute or soiffer'catial pressures. The absolutr or atutiosphtcrý%z precawie-H-ased typ patrull is used. A desirable of this t)-N is its ability to rewtai ikt ayitini cuid if the flui4 &Cal fails, by Lunga INx. case as a pressurc dtzb~n.
PRESSURE P*~1".
PK
M
C2
Y4 CCE . 110KRAIL
F~iýM
tALF
UtAI
FULL- TRAIL
vad Wi*"r,,r omba~m
0y
, IfSsR('P
K -
Lr PT11C.-Ue Traeasattmn ZEO ErEP.GIZED OE £PIýKRr Prrscee r~t1tiiiicf&e required in rcrdcr to convvlpiessu-e in ;m tehccrict%! signal !otr transmisziohaiiIWO-OI INV:V to ibr eCWkfrz in, o 0 pro-;i~z xhr j~int wiit r, visual Vildi%.aiion r. tc-f se 5chU p:hsureL. YAlthoudh NtS istandard unit;. &Prc ava ilahk., they aiv: largi I t Ltc 13-1>A~ r N t~ o ivtavy, and ob~1kkc. ?4ir~nii tarize'4.
9-17.0e
-U'C
TIhe tir~nnsiPt~ cast (.r Such a uuit i,; v)in~ ' tolc;.re 1`1V sty-skom pru~sstl,%thus C~sfli-
4available.
WRGIZEO) PoSI10 M KIOD OFENFROIFE' ENERGI ED 6*OWiICA W 2 2E THR~t-POSIT']ON VAt Vt
natiiug loss of the sys'ern when 0,:-,;cnp L.mt faiit w;s a preasurc vessel. Th/- MS sEaidalad unit. rc'quire-, a rrcn~ure tr-ansimi ts¶ibhnj devict. for pi.Ocvtivi-. plus a fuse in the tevtr the serisint, devicc fails:, *he irnnvcve4. triniataari,-d tIIIt% do not tneed snubtc~sand' fases.El
The iAc~tor valvci for nonmodutawing tonti .1 ofA ylor dclcsublystcms may be operated either wrmdlno trically. MII.-H.8775 cov--rs thuir dsgr, m~rcir-
-
N
k
t
C_.
V
'
I
1
half-tuAiiL The "trail" indirates wh.2ther dir, cylindeLt
hi;m th
h
lcc
pwosition The vjAl' it; irn the full-trril c-vndition ifth operating pn-ois ame connrcted to retirn. Fig. 9-ydtidt shchmatic zcprzsentalionl of thE variout tr;A Catiditions. Thnedvalvs mayo ýwihzer o~'.tree mapc! vl,¶ s va§y~t hav moe~fltid~b Fg. -3 shwscs -p& of dh
.4Th
two, tree an fuprij.tion vcavcs. afvU ;nay wo, U&te--, ckg'r-, the-way. or moeTh rifrttoth -ay amx- or~rt o,%th, valve. A cuRn tiretlth scvri-t.l c'uftk tK~onr for mJ
9-22
(d-PSTO AV
Iavi;i
Flilt-3.HdaAV-eCwintw one rckt'rn
sc~tiokl in &kmanifold
would be countc;d as
may bec either pilot- or dr ~j .~e1 Di'toeaion ujsually is associated with lowi1 faowt-
on,;w2v; 1 5 **
Staging, usinrz pio ',aivea. i6 :aquiuctý in orde-r to hwidle large 1. Jectively ctnd erfuicnsty is.9 8and 9-Mipwn cbmi"o bothcnt4 Mode.*A
minicauam foice kib arre qwrw for 1"ibfwav3 in catc wbmv c~miaauion of n~y uignificeaw ii pocWbk, IFoi 6,u zr4'&Ji. L~ppee spA-slsew #me. ball.jpsw type vpjvae. it t(~- mmsiaum opersuii fore ts rcqwtrwd. For the Iarprs, bwt a diý lWPe $p0KA&SmVAS a.i 40)46' iniwnm
The vaive opnwia 6mn ms =Wm!ous cm wm uvoWi -waw hammW and rewun iiids tram.SL due to r-es of 3100--psi leveb ku" *ý. rewna Wm Qantancowey. slow opersmai iaculun qs~n AM ar Oriented PMVAwily toward £-pr ~ s ms noed valveL Fig. 9-40 shows cb amimu. *Ai wre
m'r'ic
RETURN PRESSURE
~
*~
100
rOLENOID
so 60
4
OPERATE PORT
[-.
R;O
JYI-
(A)"ON FLOW" COND!TON
--
-
--
T-
OEFRAC
OPRT 'IKo OPRTI~lro
I
ýAE-
z 0fT 20
~~
'11
li
41
TME,
63G
80
insec
0O: FELOW" CONDITION
1~(B)
F~u~tM Pie'4e&~f~dVa~sl
Flzuff 9-40. Vitive %)W~daraTiq Im'
10U
The normal pr~reAS4o-eukini l.LW cbaeecwism= of the vafre wiN pvift z kAl-ea imam* if he saubsym is afkinau by EwwmIU td aw wviadventot crAvium. The level of thm IkI. pr .aw w is usch dstha ie nmwvar viLM not 6s depWdin -=a mutualduf4dm eight. Fi& "I4 Shp"s the hetum 6f this RPS udusq.4 (vinie to par. 9-2.6.1.3). W-8 aminpsswe *mcmsd is AJAIPL TR-70 AO bd tpar S. 9-2A.8.. pdiot vav Vmay be am-. i &-apam 7U~ A~pred as thwu they ac awearad to m~uara either byv -pe Low 64 tbt ammJd The swopsaewsavky
PILOT
af coe.)tIt" Mi
qwj
947.12 bsuewist R.ovicts uwew mnd as intbyua.rwolnp dsm.&u owhr pumm ovaW is masequired " They
SECIIII
Sal
-1
RETUR.
IN t~)o30"
paumw "v "~ to adtin cmuauims6 Thew mdsu. s£ix-t*j4a&-~a-ci4.rc cinpehily of f*Ul1.tidm pntwe pnwdu tham mud. ueW for kvmme cotwamisew u.j hbim Mcii.. du to pfga acoig. and is tdo paskmd mshod of uwnmd-dw apwrayift Fi 942 *gin U h ~moia am of bs* preaches
RETURN~ PRESSURE SENSING SECTION
~
frOLEN)fl
baA.1s SECTION
,Mi tkWpA fu~arpei~ w~c op"Idoli swift~t I Fisiv-941. Pt .
.
.
.
.
.
.
_
_
_ _AMCP
_
_
rSF aSysami~. l.~4
t
PlMLZRl
7W~202
_
TO MAIN SECTION CON~TROL AR~EA
(A)THE PILOT VALVE CON4FIGURATION SHOWN IS FOR POWIR RETURN OF MAIN SLClION TO THE DE-ENERGiZED POSII IO
Tho w~ntroI of the output ram is via an deetfohydraulic vulva:~ the converts the electrical cornmand in'to thec appropriate actions. The total package includcs electrical signal feedback devices, and the pivefered device is a linear variable-differentiat transformer (LVDT). The preferred clectrohy4raulic concept is the jet pipe valve, which is iolherently tolerant of contamination. Fig. 9-43 shows a schaniatic of a typical single-ýystem servo. Scpsrate servos generally are series-typc control devices, apd incorporate a position loci that isactivated when the system is not actuated or when system pressure is lost. MIL.V-27162 covers servo control valves.
[Si~lkw§:151f#-.7.14 Altawmbt l~xkv*ee Leakage should be given to leakage levels for %____Consideiation RETUN M seals in both static and operational (- LIFdynaimic situations&. ARP 1084 defints realistic in-service reqtiircriciitts for stati6. and dynamic seals.
40-
1
PRESSUIRE
)
TO MAIN SECTION CONTROL AREA
(8) TH..* PILOT VA:..VE CONFIGURATION SHOWN IS ', ýdA!1!ONfTO THEL L~ FViAPFCMttonE-EK~IRGIZLD POSIMiO# FpMm SAL.2 PeWet fi fir*~ Mh~A ft" to~48W
SAdAio Vabt
mnaj 6. mithe two-way ur one-weai. Screcc.. contdsniflatwn pcoteucon it; reqW..rd fb6oroifices with a ,J*iame~asmalm ~hw 0070in..as pea.Ji~ a MI-H.
Is~o.
Geneall. ie4 oc~wi t. f~ .u r~trcto aSin tbt. ac~uaawr cleviix. They inay bec..n caa dgc!. o p -ciAin units sucl. as the Lev. itts. Tins arranL=.-aii iri 1AUationts 0~n ;!= aisritinb~ ic~ads can cause htgh press as. In &"mi exicrna, sesit poir&aL are redcecd. Flow froma e.: 6Xhier. Qtsa&3 reqUireulient.
tor can have dte&.ructivc effects wwe-a
canaaaa.
directqy int a hydrauish host ine or eam& a &nanisr~ed
*bena
1
This should be Avoid:. A prilria 'eak w.u a ouaface restrictors is tnaa they arc felati 'ei, inacras: 4a to temperature ciiank~e; the pi.swure d, sjF for em' flouA varie* w:!h the dens -y of the fLba The nctaiý requrcaxeais for xestrlcto- .- a. en'ivcred in MIL-Ik-S440 and M11-V-25517. 9.21.13 ri Si?~1e Wtsiiatc wevas arm mr~arre for Ma C~AS. srad sasCt. 3pDWa*_..o werd requiremease11 Mr &Ma 10 aa~o :;trnmsc. into the flight commiaI aa..., Sorvoui may Pi aamsW for either singa. o; dku..
)otorhin '~atoms.
9-21 HYD)RAULIC SYSTM INSTALLATION The paragraphs that follow discuss the proper installation and support of hydraulic lines, how.s, and comaponents. The requirements also are discussed in 9-2.9A Use of Hosft aauSwich; oses and swivels miay ýe hcmy. costly, and can he a maintciiance probihmar. Thcrefore, UrW, of these items "/al be minimized. Dorirab~c alterniatives for handlinrg relative motion include coilcd tubing and fixedbody actuators. witha articulatinrg linkx between the
actuator and the control surface 0i 'ub.'sYstem furnction where applicable. 94A
NaiA4M~iac ACCO$S
n order for the- weight benefits of permtrnt:nk fittisigto be- realized, the hydraulic installation must be locatcd behuind othecr remoabwibc equipment and installations Filkcr elaioicins sh~adbe lonied icu as to~ pevnit easy &C cts MS fa doors. Diffwrenfial jprc~sut- ivdlc~iors shall
be eAhcr flu.-.n with the: &Lin or visibikc through transpair-ra sk~ii sccti-ris or nonstructural ý.ingle- or twobatitomi doors. Thi,. is neces. ary because a check ofi the pi essurc ieve: i- required ouring preflight. All saepa.-abi-. connections shals' be relatively
ac -4ssable vidoe Lhcy are employed to permit compomis .- rez~ova. afid/or remioval of lines for access to othbr cqcaipmst. t A visual a0heck efte retkmi joir fluid level condition shoLid oc possible witht,# rtmoval of access panels. I fi.. and/or bleeding is required, access to these 9-25
AMCP 706-012 i
PRESSURE 4 FEEDBACK SPRING
ELECTROHYDRAULIC JET PIPE VALVE
RETURN PRESSURE--,,,=,--" LVDT TRANSDUCER
PRESSURE__
-
PRESSURE OPERATED
"(A) DL-ENERGWZED
...
IB)ENERGIZED
Fipre 9.43. Typical Separate Servo Actuator areas shall require no more than removal of quick-
izistallation is the soft or compliant type, with damps
access doors.
or blocks close erough to the component-attach
RLS devices shall have ground-checkout capability (refer to par. 9-2.6. 1.1). Manual operating buttons may be provided, along with operating indicators on each subsystem to indicate which subsystem is shut off. Visual access to the operating indicators skhl be providud•- prefea-bly without rnoA for door opening. Access to the manual operating buttons may be through conventional access doors since their checkout will be required only periodical-
points so that excessive stresses do not develop. The installation shall be designed with the known or estimated vibration environment in mind so that sysum resonance does not oecr. Clamps and blocks may be used to advantage. D. D , it i, ,oA dcsirabc th' "hc b- -ie ,v ILI..i.. 994 discusses their use in deail.
ly,
9-2.8.4
Campo"eu
Momiti
Ceacepts
Component instaalation can be cassified as folAccess requirements are discussed in MII.-H-5440.
9-2.8.3 Hard Versus Soft Installations The tubing installations may be hard (rigidlj held in• place) or soft (compliant). Each approach may work satisfactorily; however, they cannot be mixed successfully. The use of a hard point as a fitting in a relatively noncompliant bulkhead - for instance, in cooun•tion with clamps some dist-nce away - can be disastrous. 4.t worst vibration condition, the clamps will allow enough motion of the line to that flexure at the bulkhead fitting will result in excessive bending stresea•. and fatigue failures. The preferred
9-26
lows: i. Separate. independent components 2. Inw.4rated packages with cartridge components in a common housing 3. A pressur-return manifold with 'rnount-on" components. The installation trend is to integration or manifolding. Primary reasons are tho resultant reduction in external kak points and the weight savings. An op. timum approach may be an integrated pump-iescrv,.ir package, a separate sevice center that would include all othes components exctpt the control surface actuation or suosystem operating device. In any
even, each helicopter desin must be a
.d
indi-
vidually in conducting the system installaition tradcoff.
S(:Componws Ssumch •. '•
SNatural
S.LJ !; M isel nv w '"- lbtk e Co ukka~k oli Components roluirL" frequent maintenance shall be accessble. Such units as flukiddrivcn air comprcssomf pumps. and filters must have unusually good &c in order to allow adequate se~rvicing. that require firquent maintenance, as filters and pricunimfic chemical driems. AW1 be mouted rigidly in onrde to avoid damage during acrviciall, dirt-collec.ng area, such as brake valves
vmde eazc hocepit fldor. ss be onsidefd in planvialsysi.m irdallation. :rati-ue A typical exampic of improthr instaeltion is routing of line through a walla-wcl area, where leaking 2ir car causu strumming of both normal and Cmpoenc t res. Attachment of li to n e panel thal may vibrate will have a similar Wect. Proximity to a common fire hzard aflo must e avoided. Separaaion of kaltnbea accomplishlo aater consideration Cof mpll ,vionemn.'t-s hredquir freluate n
The
gn ofactuator vnlving should include pro-
vizinna to prevent fluids being trappedl at the ends of tk• main controi-valve spool. Actuator pistons shallnot bottom against internal actuator stops during autopilot (stability augme~nta tion) input, coupkle with the :!xtrmes of manual input and rigging and manufacturing tolcranca. Erratic stability augmentation inputs during taxiing can cause "ccsive loads due to hammhering aginat stolp.. T]his condition "dd be taken into account durin testing, as well as in dezign. Internal hydraulic stops in cylinders "~ incorporate snubbing. or AdJ be sta~the~ned suffi•ciently to prevcni fatiguve loading failures (Fig. 9-45). Rigging
aTone should nota conrdevral t in orlder to picrnt lons oni. However, tra desgner should rrmai controls to stick-stp limits to ovoid actuator bottomaig. To prrvent blow-bypTeflon piston seals should be avoion in an actuator that also must operate pnuumatically. Piston rings shan m be used on piston heads. Appropriatesails should be onsidnsdu for longstroke, chanat earge-bore cylinder ofactuator roThinsand subet applications iucant dartiagcc too
muThe rsigil• rfaclure avoid instaleations which neab l enaina mount canto1 drop the enginein on cr an piig:f lines or components and thereby increasepnNTE the simnibility of losing thenuT helicopter. --- "
•9-.9
()l
D He
DESIGN CRITERIA The MISCELLANEOUS designershalaodisaltosi.hc sas laIndustry experience s ith previous hydraulicn sysvte-hn and has rhveallbumane mompontnst cancous dfirn asprts thal halube a consideredf. Th ofragraphs that follow phaarnt some of thea cont
Do T
.1
o
iiscnl-
For protection against dynamic shaft seal failures, the "~grir should consider use of dual seals. with a return vent incorporatadbetwhen the seals (Fig. 9• 44).Figure Thp design of the electrical system with regard to cietrohydraulic: servo valves shall avoid
9-04. DualI Seals With Retur
Veat
dither
nsgnalr cause ehar valve osillation and actuator dynamic si wear. Pronetion against electromagnetic interference susceptibility shallbe rcequired. Linkarg pivots on flight control actuaturse con2in- of friction-rsld journal bushings withDoore Ftolrparnoe, iro crtical yhimming in ordseato of an dthe lignMero necossary for frue ofiration. s whall•
bearingand pivots,op•rratin with provisions overtorar t prevent uid in d systemto return
quirn pnTherd ligknls.
in order to prevent binding of control-valvr oscilat
Fae 9c4u. Hyviavlve
for lacetaal Sds 9-27
--
--
-
'"
".
AMCP 70&20 Internally threaded arcuator endcaps utilizing AND 10050 boss scaling design. as shown in Fig. 9•4(B), should be avoided. Leakage can be caused by squareness of threads to boss surfice and torquing problems. In addition. end caps that are threaded internally into a cylinder barrel and locked with a jam nut arc subject to barrel stretch under pmsure. This can result in leakage or loosening of the jam nut. Sufficient material thickness in the cap ama is required. Viscous dampers sh•l be self-servicing from Ssystem return fluid. for servicing of hydraufittings grease be: of Zerk hel•kl avoided. ccwmponets lie Use
piston diamet sAdl be provided for piston and rod bearing amas a the fully extended position (afe Fig, 9-47). Safcty-wirinn of piston head raining nuts to a piston bead that can rotate on the piston shaft is not acceptable as a locking method. Use of -vkting devices such as the NAS 559 keys as shown in Fig. 948 siuld be considered.
In order to minimize binding and seal wear from actuator side loading, a minimum overlap of one
ir, .wder to incrase the presur required for a given brake torque is a suitable corrective action. The brake control valve input ihaf. s/U have adoquate beanag surface area. and should not be saijected to sde loakdn from hosms, ctc.. so as to 1asure saooth brake valve operation and full release.
%L91 Brake Design Exac.veut
p).Ruli
brake to e st wheel ebrak o effctoiv w dbrkphoar
oe."'t
BEARING OVERLAPONE PISTON DIAMETER I
_1
(A)CORRECT Fiswe 9-47. bearing Oveirla
(B) INCORRECT Figure 9-46. Avid Internally Threaded End (aqm 9-28
LOCKING KEY Figure 9-48. Plson Head Retaining Not Locking Key
AMCP 706-202 9-2.9.3 Coatrol System Design Flight control system assemblies shall have adequate clearance guards, or otherwise be protecked, to afford maximum protection against jamrming by foreign objects. The designer shall avoid routing flight-control linkage through areas in which its removal is required in order to replace the engine, Cable tension-retaining devices shall be considered as a means of preventing control cable tension changes. Overtorquing of control system bolts shall noi result in increased friction during operation. The use of special bolts shall bc avoided, 9-2.9.4 Electrical Design Electrical connectioni to hydraulic components shall hav/e a mechanical strength iequiremen consistent with maintenance handling requirements Wires should be buried in the installation if possible. To simplify troubleshooting and component replacement, hydraulic or pneumatic components incorporating an electrical function shall have integral electrical connectors for removal and replacement. Puiili 5 pi.•aun a•wt ni-o .c..ir. a hl;ghcr I.•., cure than can be withstood by electrical insulation. The compounds also must be compatible with subscquent processes applied to the assembly d,,ring manufacture, such as welding or baking for epoxy cure or strain relief. Proper manufacturing of electrical connectors rquires that only the wire should enter the soldering connection, The first layer of insulation of the wire should enter the potting compound so as to provide
moisture kakage protection (Fig. 9-49). The braided
insulation should be clamped adequately a the connector inlet, and should not enter the potting compound in such a manner as to provide a leakage path.
Higher quality electrical parts should be used in place of MS parts in critical applications where failure creates a high probability of catastrophic
CLAMP
effect. Detailed failure analyses of electronic and electrical circuits are required in order to determine where use of such higher priced parts is justified. A positive fix is required in order to prevent runaway trim actuators. One possibility is stepped motor operation. Two electrical actuators in parallel, with braking when dc-energized, must have indcpendent electrical inputs so that the fitst actuator to complete it%stroke can be de-energized and braked. .9.$ Filter Design Filters shall be installed in the pump drain line prior to its eotry into the oil coolers. (Installation downstream of the coolers will allow trapped pump particles in the cooler to recontaminate a replace filter element.) Proper flushing of a cooler is important. For T-valve inst'liations, central filtration should be used in order to avoid differential flow as the individual filter pressure differential changes (Fig. 950). All restrictors with hole sizes of under 0.070 in. should incorporate filters.
PRESSURE
6LE
HLTFR PRESSURE FILTER
INSULATION
WIRE BRAID PiWr
POTTING
(B CORRECT
-" . Properly Aembl• Electrical Couaectow
Fiture 9-50. T-talhe (eeiaul Filtrvaim
r
'U
~7W2.V2
Test filtration should not exceed that of the component in actual use. 9-2.9.6 Fittings Design Hydraulic fittings, such as AN g33 universal hulkhead types- when installed in a valve port, can result in internal valve interference or restricted fluid flow. Component port design and fitting selection shall be such that interference cannot occur. Pump fittings - and suppressors, if incorporated Sshall be torqued to maxinmum allowable values in order to prevent loosening and subsequent loss of fluid. Use of an acceptable locking device is advisable for any large fitting in high-pressure application. The use of pipe plugs for external sealing of drilled passages can lead to internal stresses in component housings, resulting in cracks. The Lee plug produces less stress concentration and should be considered. 9-2.9.7 Cage apd Indicator Design It ageneraly gndit d Design tlet It generally is good design to include fuses as well as snubbers at the upstream ends of the lines leading to gages and indicators. An alternative is to install the pressure sensors into fittings in the system line instead of in an annendant line. Helimcil sennine elements are recommendod instead of 5ourdon tube types. Gages with Bourdon tube or rack and pinion gearing shalltransmitter be avoidedshould for use pressure toindigcators. The be as conneted the
. .
SGage,
include suction and clemznd requirements based upon the condition of the application.
buildup of case pressure if the sensing element leaks. The vent hclt should be covered with tape to prevent corrosion.
common routing can result in priming problems, with a momentary interruption of inlet flow resulting if reservoir pressurization is low.
9-2.9.8
cavitating a bootstrap reservoir, the auxiliaiy pump
Hose Design
Realistic design and testi'g of devices such as brake control valves must take hose side loads into 4,.count. Hose routing or sizing shall prevent cross-connection at actuators. However, deviations from normal hydraulic practices, such as use of return hoses that arc smaller than pressure hoses, shall be avoided because such deviations can have an adverse effect.
Sof
9-2.9.9 Pump Desiga Insuring compatibility of the pump with the system requires determination of the effects of low inlet pressure, high case drain (or bypass) back pressure, and the interaction of the two on the internal balance the pump. Back pressuie also can cause reduced 9-30
I'
Pump cavitation will result if reservoir pressurization is not sufficient to accelerate the fluid in the suc- tion line to a flow rate compatible with pump displacement. This condition is likely to be a more cri- "tical design condition than is the steady-state flow requirebnent. Qualification testing of pumps, larly those for use in power-control systems,particushould
system with flexible hose or with a tube having sufficient bend to absorb vibrations, cases shall incorporate a vent hole to preven"
,i-
cooling flow, leading to shortened pump life. Compatibility determination includes analysis of the nature of the contamination generation properties of the pump; sufficient filters must be used to keep back pressure low within a reasonable cleaning schedule while maintaining a clean fluid supply for the pump. Two-pump system design shall consider large, nonbypass filters in the drain line of each pump. Should bypiss-type filters be used to insure low pump case pressures, the flow shall be routed through a second, larger return filter. Pulsations resulting from pump ripple, which may be intensified by system resonance, can be determined by oscilliscope scanning of the pressure through the range of operation. Peak pulsations shall be kept below +150 psi (300 psi total). Pressure pickups must be in the line (not on appendages), and shall be located at the pump and, at least, at the first downstream component. The optimum design furnishes some elasticity to the system at the pump outport. Short, dead-ended lines near the pump require particularly close scrutiny, and should be avoided.
Connecting two or more pump systems from a
ii the normai sysiem pump is pumping air and
can be affected adversely by the low bootstrap supply pressure, and may not prime. A check-off accumulator may be required in order to maintain bootstrap pressure with losh of normal system pressure. Centrifugal pump operation with an outlet flow blockage can result in overheat, thus causing seal or case (structural) failures. Therefore. bypass flow for cooling shall be provided. Pump testing shall include realistic case drain system chara.-teristics. The internal leakage of a hydreulic pump is necessa.-y for lubrication and cooling of the pump mechanisms. However, the subsystems into which the flow is discharged vary among helicopter designs. The designer shall spocify the case drain system characteristics to insure adequate
AM housing strength, shaft meal capability, and pressure conditions during the pump qualification testing. Water hammer limiting is discussed ;n par. 9-4.2,
AMCP 706-201. 9-2.9.10 Resenolr 1[)Ip Bootstrap reservoir design shall incorporate sufficient piston force, in a static, no-pressure condition, to facilitate reservoir servicing and bleeding. Reserv.iru shall be designed with the air bleed vent high and the suction outlet low. The overboard relief flow capability shadl be sufficient to prevent reservo3ir damage during improper or emergency operations, such as system operation with an ovcrfilled reservoir or overfilling during reservoir servicing. The designer should avoid the connection of two or more drain or vent lines together, to a common overboard vent, where back pressure can cause back flow through the second vent system. With hydraulic power present in one system only, high rates of motion in large, tandem actuators can pump the fluid from the unpowered section back to return system without recovering equal fluid from "thepreuture side of the unpressurized system. Unless exist to dimp the rttur-ed fluid at a low pressure, damage to the reservoir and other lowpressure nomponents can occur. An alternative to dumping is to equip ground test carts with multiple connections so that both systems may be pressurized simultaneously during checkout. Test reservoirs shall be representative of the actua! .system reservoir. This will allow viscosity, fluid ternpereture, fluid settling, and flud aeration test conditions to be realistic,
Sthe
)provisions 17-
Chock valves shalbeis t d squate ~~~~Check valves shall be installed ir subsystem return
0•2
lines so as to prevent back pressures from high return flows from acting upon cylinder locks, diff-rential cylinder areas, and return cavities in components that may fail under repeated return transient pressures. The miniature check valve should be used for this application since it has a faster response time than the standard AN type. In addition, balanced areas for lock devices are recom nended. Check valves shall be install d in the pressure lines
of subsystems where airloads can cause a flow revetsal when system pressures are reduced because of an operational demand upon the system. A relief check valve shall be considered if overloading can occur at
7
202
corrected by installing a fast-acting miniature check valve in the return line. Soft seals or poppets that depend upon assembly
compression to prevent secondary leak paths around the material may leak duc to distortion or compression under operating pressures. This can be pieventcd by incorporating static 0 rings to protect secondary leak paths (Fig. 9-51). The designer shall avoid use of self-locking nuts to hold spring-loaded adjustments. Part concentricity shall not be dependent upon thread concentricity. Flow paths within valves shall be considered. Indexing radial holes in spools or placement of springs can affect flow paths. Poppets with flutes can rotatc with flow, and thus may be desirable in some designs. Inadequately designed spring guides may allow spring lands to cock spools. Center point loading at both spring ends is rec'immended. Reduction of load upon springs that are heat-soaked while loaded must be a dcsign consideration. The designer should avoid tension and plated springs. For springs immersed in fluid, etc., 17-4 PH spring material should be used (refer to MIL-HDBK-5). Low-operating force valves, such as solenoid-vaive pilot sections, should be designed with poppets since spool valves are subject to sticking from contamination (silting, etc.). Vent holes between two seals of differential areas, such as arc incorporated ,. return-line dampers and inline balancei relief valve designs, shall be multiple or indexed te insure that a leakage trap does not occur. Urit malfunction may result if the vent chamber betomes filled with fluid that is not readily dischargable through a small vent hole.
eDirect-operated solenoid valves shall have adereturn spring force to overcome silting act.on (hpsern) (chip shearing).
POTENTIAL LEA K PATH
r
II
-
ADD SEAL
hg) ih speeds. .
-
Reguleaed pressure can be affected by transient
baick pressuues at the valve return port This can be
Figure 9-51. Secondary Leak Path Seal 9-31
AW
70&202
_____________________
Split-coil (holding) solenoids should be avoided in valve desijln. Upreliab,lc operation, caused by starting-coil switch aettings and malfunctions, results in overheated solenoids. This problem usually overrides such desirable features as the lower weight of split-coil designs. Servo-valve-cover deflections or shocks can cause valve manfunctions if the cover ihmounted upon the semro valve motor instead of the main valve body. Adequate clearance between cover and motor should be provided in order to minimize effects of minordents. Covers should be of ungged design. Hydraulic presure surges du- to valve spool shift may be prevented by the use of slow valving (adequate dead band) or electric time delay. In thr case of eloa.trically operated valves, consideration shall be given to valve positioning, and to the effect upon the system should electrical power be applied inadvertently to two electrical inputs doe to shorting, etc. Relays can insure predictable operation in this abnormal situationpower in hyd abnormaulc eme uen.y odesign Hydraulic emergeno cy oveboard dump vaivpssal have sufficient flow cap&city to avoid back pressure buildup th-t, in some system. dsigns, can divert fluid to the reservoi. through selector valves and check valves. Directing fluid to the reservoir may result in reservoir overpressurization and failures under certaD conditians. SDifferential arca vent sc~l wear ad leakage can occur fron plunger fnation fiuiring normal system pressure fluctuations. Long-li;e seals, or valve designs incorporating little or no plunger motion during normal system pressure fluctuation, are desirable, In the case of half.trail valves, internal leakage
"•
when in the neutral half-trail position should be evaluated for its effects upon subsystem operation. When leakage from pressure to blocked cylinder port is greater than leakage from block cylinder port to return, pressure buildup in the blocked circuit can occur, resulting in unwanted motion or loads in the blocked circuit. Excessive leakage from the blocked cylinder port toloads. return can result in unwanted motion from external
xteral High-pressurefrom tests with returnlads.The ports capped can
-p
result in overpressurization of components. The return proof pressure rquirement should be The patible with system operating pressure, -r
9-2.9.12 Lubrcation Experience indicates that graphite-loaded grease tends to dry up in high-temperature antifriction bearing applications, leaving a residue of hard graphitc that interferes with proper bearing function. 9-32
All critical joints .ala/l be labri-ated and protected. L The lubricant must be compatible with oiling. Lubrication of mechanisms that Are located in a high-temperature area can result in jemmlWg as a result of burned oil carbon. Left- and right-hand component lubrication fittings shall be multiple, or shall be located so as to be accssible. Long lubrication paths result in frozen irease and blocked fittings. Unclamped monoball bearirigs must have two grease fittings in order to insurv proper lubrication on both ID and OD. 9-3 PNEUMATIC SYSTEMS 9.3.1 PNEUMATIC SVSTEM DESIGN Te decision to us a compressed gas rather than e pressurized liquid as a working medium in a fluid poer lisysemas a during tediminafy ontrol system is made during the preliminary phase. The various tradc-offt to be considered in making the choice are presented in Chapter 9, AMCP 7 design paragraphs that describe a pneumaticforfollow the detail 7-201, Theonsiderations system. 9-3.1.1 Systo i Analysis The design and aialysis of a pneumatic system become considec'ably involved when nonlinearities are considered. The derivation of a mathematical model describing the physical phenomena of cornprssible fluid through a system -- where the fluid through restrictiorfcepansionu,changes in passes dire, tion, etc. -- proves difficult and results in cume. •_e.uat.o. . For example, the flow within the system, or the pressure drop, will vary between the extremes ofadiabatic Pfow (no heat transfer) and isothermal flow. The basic formulas for adiabatic and isothermal flow are given in par. 9-5.5, AMCP 706-201. The formula for calculating the mrximum ms flow of ai: within a system is pr-tended in the same paragraph.
calculation of lwtruhnzls flow through novIxcs, orifices Tecluaino rtws piping, valves, and fittings may be simplified by use of charts and graphs for expansion factors, orifice coefficients, iftical pressure ratioc for zzles, and relative roughness and friction factors Zoi p .hng and tubing. Familiarity with the simplified c4uations of Refs. 2 and 4, and use of the tables, graphs, and charts cont'iined therein, will allow a good analysis of an entire system or component to be made. rhe venfication of the design through actual operational test of the system or component performance is the designer's ultiinatc goal. .
-o
706.202
___MCP
4 v
C S
To aid in the selection of particular components, a list of commonly used comnponents, and an operstional description of each. is provided in par. 9-3.2. Refs. I through 4 arc additional sources of information regarding design and analysis of pneumatic cornponcnts aaud systems.
An APU may be used for providing emergency rneumatic power. The APU shiould be designed to use stored c'mpress,-d air for starting and then to provide a limited amount of power for the essential subsystems of the helicopter.
9-3.1.2 System Redundancy Ali pneumatically operated services that are essential to safety in flight or landing shall be provided with emergency devices per MIL-P-5518. The emergency systems must be completcey independent of the main system up to, but not necessarily including, the actuating cylinder or motor. These emergency systems should be designed to be actuated onty by cornpressed air, direct mechanical connection, cltctromechanical units, gravity, or combinations of these, Where dual pneumatic lines are used to providc emergency operation of a mechanism, the normal and emergency lines shcll be separated by as great a distance as is practicable, so that the possibility of both lines being ruptured by a single projectile is remote. Where shuttle valves are necessary in order to connect .he normal and emergency systems to an
9-3.2 COMPONENT DESIGN 9-3.2.1 Air Compresirs The air compressor maintains the pneumatic system pressurization during flight. It can be driven by direct drive from the helico'pter engine gearbox, by an electric motor, or more commonly, by a hydraulic motor powered by the utilit) hydraulic system. Compressor operation usually is controlled by a manifold pressure sensing switch, with the compressor cutting in when system pressure drops to a preset minimum and cutting out at a prescet maximum. Compressors can be classified into two basic groups: positive displacement, and dynamic, or nonpositive, displacement.
u in w..t, ..
u,....
,w ....
- •
'-.-
i,.:.t
...
*k.
cylinder. The emergency line from tPe shuttle valve should be vented to the atmosphere When not in use. When an air bmttle is used as an emergency backup energy source, a standard pressure gage shasl be installed to allow maintenance personnel to check the pressure. The air bottle should be located so as to produce a minimum length of line between it and the shuttle valve.
9-3.2.1.1 Positive Displacement In this type, pressure is increased by confining a o in n prnoreL*Wsivelv dirninishini svace. There are a
on ... a_ . -
..
number of different arrangements, among them- the axial piston and the rotary. In the axial system a piston moving within a cylinder (Fig. 9-52) alternately traps and compresses the gas. This is the most widely used type, and sizes range from less than I hp to 5000 hp. Good part-load efficiency makes this ty;3e most acceptable where wide variations in capaci.vy are required.
INLtT
OUTLET
,9 "Figure 9-52. Axial Piston Compressor 9-33
Wtik
......
....
.
. . .
.....
... .
In the rotary system, the rotating motion of single or mating elements compreLses the gas. Major types of rotary compressors are sliding vane (Fig. 9-53), "lobed-rotor, liquid piston, and helical, 9-3.2.1.2 Dynamic Displacement In this type compressor, a high-speed, rotating ecment imparts velocity to the gas. This velocity is converted into a pressure rise in the compressor volute or other diverging paMsageways. There are two main arrangements, centrifugal and axial flow. Centrifugal compressors have an impeller similar to a centrifugal pump. Impellers can be arranged singly, or in multiple units for higher discharge pressures. At a constant speed, a centrifugal compre3sor delivers nearly constant discharge pressure over a considerable range of inlet capacities. Axial-flow compreassrs move air parallel to the rotor axi3. They are made in single- or multiple-stage versions. In the latter, matching stator blades redirect the flow of air to the proper entrance angle for sueceeding rotating blades. Generally, axial comarressors are used for ultra-high capacity. However, there are -,J•
many gri.' ..... •
annhlictinnt for qmaller unit., r ......... ............
9-3.2.2 Compressed Air Supply System Selection and Operation There is overlap in the performance of different compressor types, and sometimes several can be used for any given service. Naitowing the choices is a proces of considering such factors as space and weight limitations, power ranges, and capacity ranges of the
the air inlet cap on the first-stage cyiinad head passe.n through a filter to allow mmoval of any pwrticsa that might damage the internal components of the assembly. The fittered air is drawn through the intake valve into the first-stage cylinder by means of the saction created on the downward, or intake, stroke of the first-stage piston. On the upward (compression) stroke, the intake valve is forced shut by the increasing pressure., and the spning-loaded discharge valve is forced open when the presaure reaches a pivdetermined value, The comprese• air is directed thrc'lgh the discharge valve and into the first intercooler connected to the first-stage head, where the heat created durir.g compression is dissipated through forced convection by the airflow from the fan directed over the intercoolers. The flow of cooled, compressed air next passes through the first-stage relief valve - connected between the first intercooler and through the inlet port in the second-stage cylinder head - and then into the second-stage cylinder during the downward (intake) stroke of the second-stage piston. The operation inside the three subsequent stages is identical to that of the first stage. e • . t ,' Moisture s-parators (Fig. 9-54) are used in 'zonjunction with a chemical drier as the dehydration equipment of a high-pressure pneumatic system. These units, wor king together, deliver dry air having 9,-3.2.3
,.AToD
different types.
W
R VALVE• ,WLATING .
In a typical helicopter compressor of the rmciproeating type, the air entering the compressor through
PES.',
. ,,a
r
V
RI•LIEF
ENTRAPPED
INLET
I
ACrUAlED BY
W OPERATING
Figure 9-53. Sliding Vane Compressor 9-34
-
OVIRINOWq
CO----) %
\
(8) BLOW UDOW
Figulre 9-54. Meoistr Separator Imcerporathg Systsi Fressm re Replathig Valve
)
-'when
-
~~AWC 706=~
a free air dewpoint of -65'F by tra.pping and coilectiag droplets or moisture that literally have been squeezed out of the air during compression. In a symtm preasure of 2W) atmospheres, the separator red~wes the free air dewpoint to -. 15*1. Collected moisture is drained from the separator either by a dump mechanism that operates automatically when system pressure drops upttream of the separator. or by a mczhanisi *;hatdischarges the moisture at frequent intervals while the compressor is supplying air to the system. Th; air compressor incorporates Pbleed valve that #1lows the separator to blow down automrati'ally the compressor stops running. During this operation, the compressor and the interconnc:tins lines to the separatoo. also arc blown down. The moisture separato. includes a heating unit that prevets the ac~cumulated water from frcezing. Also included is a Aafety disk that protects the separator from the effects of overpresaurization. A backpressure valve is used directly after thc moisture separator in order to build up pressuit in the moisture-separation chamber before it cat. build up int downstream com~ponents. This vaivt ii..ura timmoisture separation whent the compressor running. The vitlve either is integral to the separator or is installed separately. Valves are available with various back-presarc settings.
)mediate
-~starts
9-3.4
Deydiaarspulp,
The oasivaor-rmovl typs eluipentare mechanical and chemical. Mec~i-nical dehydration usually 'des a refrixeration cooling process that iowcrt the' air temperatur bCOWtbcrvdp3nt. The CjndeqtLW water then is collected and eliminated. The limiting factor is the temperature to which the air can be lowered. This type usually is not found on airborne systems due to the weight penalty of the refrigeration cquip~nent. Chemiical dehydration normnally is uswd in conjunction with a moisture separator in order to providc maximum efficiency. The chemical drier is placed immedittely afer the moisture separator (refer to par. 9-3.2.3). Trhe pneumatic system bcyornd this dehydration equipment thus operates with~ dry air, reducing the possibility of freezing in lines or17 Although they are called cheraical Vcomponents. driers, these units reduce moisture content by the procx~s of absorption, and no chemical change takes place. Each unit consists of two parts: (1) a metal housing that acts as the pressure c~ontainer, and (2) a replambcal carrtridge containir g the drying agent. The life of a cartridge will depe. A not only upon the
rate of airflow, but also upon the ambient and inlet air temperatures and the moisture con-ent. If thIe chemical drier is to be serviced manually, a filter sW~ be included at the outlet port so as to prevent downstream~ mugration of particle& of thr. me placement compound. Some means of indicating when the compound is no longer removing waiier vapoi efffectively sA1./ be included. In the manually reactivated types. eae of compound remov.41 for replacemcnt or reactivation should be considered. The compound should be in cAttridge or c~apsule form to prevent spilling. 10.2.5 Fllteii For long life and trouble-fret performance of system components, the air should be kept as clean as possible. Fig. 9-55 shows some filter confligutrations. There are many sources of cont&minauion: the air itself may be contaminated by dirt from the atmosphere, or from system hose connections and other transfer devices. Pa.tidcls from worn system components are significant sources of contamination. All filter media act to varying degrees as both wkp"tb afi~d Surfdc iic Howev~er, !he rma.y-m classified on the basis of the predominant type of Fril tration provided. Depth media depend upon long, tortuous flow paths to remove ccrtaminants. Examples of these media include Paper, Vclllos,10. Wet. glass fiber, wood and sintered powder. In fiber filte-rs, variations in thickness, depsity, and fiber diaimrcter are comnbined to produce nominal filtration of 0.5 to 10D mirn.Aslt tig ar io o5 ie h mirn.Asltrtngvayfo2to0tmcth .
F7i
r4
777
(A)T-TYPE
(S'POT-TYPE
l, i~i.,L
(C NIETY
()ILN
YE()VTP
D)-YP
Figure 9-55. Filter Housiag Designs 9-35
AWC 7(W=( nominal rat A. sintere powders of metal, ceamic. of plaistic provid depth filtration with nominal ratings of 2 to 65 microns and absolute ratings of 13 to lOD microns. Media migration is t~te primary disadvantage of sinteired powder, but this can be avoided by proper manufacturing techniques. Surface media remove contaminants by means of s surface that contains fairly uniform orifices. Thus, the contaminants are retained on the media surface, In wire mesh types, small, uniform-dianreter wi~res, woven into a Dutch-twill or square pattern, provide nombial filtration ratings of 2 to 100 r~aicrons and absolute ratings from 12 to 200 microns. W.re mesh has good strength and is frcc from media naigrmnion. Dirt capecity per unit area is low, but the thinness of the mesh permits use of multiple layers.
"9-.2.6 Valves Valves include any device that stops, starts, or otherwise regulates the flow of a fluid by means of a movable element that opens or obstructs a flow passage. The most commonly used valves in any airborne pneumatic systun are described in the para-
no damping can be incorporated irto the v'echanism? of the ball check valve, and the chattering tendency ca -not be eliminated. Therefore, bell check valves are not recommcrnded for applications where chattering is unacceptable. In cone check valvies, the ball is replaoed by a sliding element with a conical sedtialg surfsc at one end, This surface seats against a circular sharp a*r or another conical surface. Cone check valves generally have less pressure drop fc'r a gi's'n size than do ball check valves and have less tendency to chatter 'L-ause of the guided movement and resultan: damping of the valving element. Cone c"ec valves are susceptible to dirt ini the seating are* and between the piston and body; this can cause cocking. with a resultant leakage between the piston and the body seating area. Cone check valves generally are used in the same types of aipplications as are ball check valves. Howe,.er, cone types can be uscC to produce a reduced pressure drop in a valve of given size, and also can be used in applications whent the tendency to chatter cannot be toloiatod.
graphs that follow-
______
9-3.2.6.1 Check Valves
twrnp^f arissg.r¶mnt riO~ffamn~oo ni
NO
UNRESTRICTED
The primary function of a check valve i6 to prevent flow reversal. Check valves pass air freely in one direction and, if pressure reverses, close quickly to stop flow in the other direction. Flow reversal in fluid systems may be programmed as a normal occurrence, or may be caused by accidents or failures. Accidental flow reversal must be halted promptly and effectively. If this is not done, accumulators may be overprc-ssurized, rotating equipment may overspeed, or other -. p,,.r
FLOW
F FLOW
(A)SALL CHECK VALVE
U4RESTRiCTLD
FLOW
NO
811
FLOW
('hc.ck
arc automatic in their opetation, with their valvinig elements being acuiv~ated by the foices of the following media. Four types of check valves areN shown in Fig. 9-56. In the ball check valve, a hardened ball serves as
_______________
J8) CONE CHECK VALVE
URSRCE..,RSRCE FNETE) LOW
FLOWE
the closure element and is spring-loaded against a circular, conical, or sphefical scat. Flow forcez; lift the balloffthesca an aginsttheloaingsprng.(C)
Because the flow must proceed around the beill, this type of valve shows more of a tendency toward turbulence and pressure drop than do other typc4. During normal operation. the ball rotates slightly on the retaining spring, thus allowing eve.i wear on the ball and the valve seat and minimizirg the effects of contanination. Because of inherent simplicity and low cost, ball check valves are used frequently in applications involving small ;ine diameters. where pressure drop is not of particular concern. Practically 9-36
RESTRIL~iwnrCHEt.K VALVE
UNRESTRICTE.D FLW
__
*0 (D) POPPET CHECK VALVE
Flgiere 11-36. Check Valve
NO 'FO
.
) A variation of the cone check valve is the restriction check valve. In this type, full flow is allowed in de forward direction, while a restricted flow is obtaimed in the reverse direction by meaps of a small orilice in the conical seating element. This type of action cannot be achieved in a ball check valve. Poppet check valves consist of a mushroom .shaped poppm, with the stem closely guided in the valve body and the head sealing against a flat or ta-3ered circular scat. In this valve, flow forces in the forward direction lift the head of the poppet off the seat, and flow proceeds through the stem of the poppet, around the head. and through the body of the va: . In general, poppet-type check valves have less pressure drop for a given fluw rate than do either cone or ball types. Poppet chock valves can be desllned so as to eliminate any tendency toward chatter or hanmenring by the incorporation of damping chambers :n the valve. Because of the close clesiances between the poppet stem and the valve body, contamination can cause sticking aid leakage. Popoet check valves require more parts than ball or cone types, and. therefore, usually are more costly. They arc used most commonly in applications where it is desirablc to improve flow characteristics,
At
The primary usv of relief valves is to controi fluid pressure in a tank or system by cischargiug excess flow to an area of lower pressure. A relief valve is a prsure-relieving device that opens automatically when a predetermined pressure is reached. Relief valves may have a full opening "pop" action, or may open in proportion to overpressure. Valves that open rapidly to full flow generally arc referred to as safety
4
valves or pop valves, and are considered a special fomof reie alve. A relief vilve consis.s of a valve body, a reference load, and a closure that serves as a control element and seat (Fig. 9-57). Ihe reference load is linked to the closure, and opposes the pressure buildup in the tank or system. The magnitude of the load determines the relief preusare setting. As the internal system premsur. increases to neadly the relief pressure neoessary to balance the reference load, leakapc usually begins. When the internal pressure reaches the relief pressure level, the valvc opens and discharges the upstream air. As internal pressure decreases below the set pressure, the refircnce load opposing the pressure force closes the valve. A relief vJvc is considered to have good operating I characteristics when the pressure for rated flow and "eseat closely approaches the cracking precure. The cracking pressure is the relief pressure setting of the
AMCP 706-202 valve, defined as the pressure where lAkage flow reaches sonme specified value. The cracking pressure always is set below the allowable working pressure of the tank or system, and commonly is not more than 110% of normal operating pressure. The rated capac;ty usually is established for flows at pressures 10% greater than the pressi,'• setting of the relief valve. The rseat pressure is some value below the cracking pressur., dependir4 upon the closure configuration; a restat pressure of 95% of cracking paivsure is common. Relief valves may be either direct-acting or piloted. Diret-acting valves can be either of the conventional type, where the control element moves ielative to the itat, or of the inverted type, where the smat moves relative to the control element. For airborne applicatione, the aelief valve body is desigi~ed for minimum weight, consistent with pressure rating, and for passage of high flows with minimum pressure loss. Lightweight construction materials, such as aluminum, arc used extensively. To achieve minimum prtssure drop, some manufacturers use a venturi design in the discharge side, while other- enlarge the outlet port even to the extent of using larger connections. TFL " t-• 8 •. .L. -. .. .. . ; o... . pressure buildup until relief pressure is reached. The most common element usod to establish the load is a compession spring. Weights could accomplish the same purpose, but seldom are used. Is
cr
___.I
I J
m[1
F.) z.-1 ct.•
. 4,
__z•
r
"ii!ŽI"
+ (A) BALL
(B)CONICAL-
POPPET
a--
*
(C) V-POPPET
(D) PISTON
Figure 9-57. Relief Vaives 9-37
TI
*
The valving unit is composed of a seat and a control cie'ne.a. The seat may be flst, sphierical, or conIce. inl shapt. wit its configuration determining the sealing And opening characteristics of the relief valve to a large degree. There aie four commonly used control element daiglns: ball, conical-poppet, V-poppet, and piston. The ball control element isused extensively in both quack-opening and proportional relief valves because of its simplicity, low manufacturing coat, and inhecrerit self-algning capability as it rescasts. When used in a quick-opening relief velve (safety valve), the ball tenda; to chatter when discharging fluid. It is limited to small valve sizes, and has a short life cycle, Like the ball control, the conical-poppct may be used in both quick-opening and proportional relief valves. Conical-poppets lend themselves to larger port sizos, but require closer tolerances on the seAling surfaces of the poppet and seat. The control element stern must be guided in order to obtain alignment between the poppet and seat. When the valve. isopen, however, the instide surface of the guide is exposed to the fluid. If ags is discharvcd. cooling dur~no expansiovii may Meult in an ice ulupoui h ieta will prevncrt the valve from closing. This type is quieter in operation than the ball, due to frictonal damping induced by the guide. The V-poppet element isused only in safety valves. As soun it. the valve starts to open,* the fluid - by changing momentum due to the V-design - exerts a greater force against the poppet, causing it to pop open for full flow. The poppet uses only its inner cone for a sealing surfacc. Like the conical plug. the poppet stemn isguided. Precision machining is required in order tO obtain a -ar9!C poppnet-ea and alignment, A piston sometimes is used in relief valves for closed air systems. The piston offers no positive scating surface to prevent leakage, depcnaing primarily upon close tolerances. Valve opening is proportional to the overpresaure. Pistons arc used comnmonly "s the second stnep in pilot-operated relief valves, rxther than in single-aage valves,
".U.3.2 Preuwe-edsciug Valves The most practical comiponents for maintaining secondary lower pressures in a pneumatic system arec pressure-reducing valves. These are normally open, two-way valves that &:rise downstream pressurc in order to close. Tlhere are two types, direct-acting and pilot-operated, Direct-acting valves are usually of sliding spool design. Air flows from the high-presure inlet to the low-pressure outlet. An adjustablc spring holds the 9-38
poppet or spool open, and reduced pressure aims to close thc valve. When the valve is closed. a small quantity of air bleeds from the low-pressure side of the valve through the sprnpg chamber to the ammmiphere, usually through a fixed or adjuatable orifice in the 4pool or body. The bleedofl prevents dowastream prmsure from increasing above the valve satting because of spool leakage when the valv4 closes. The spring chamber always is drained to 4.%e atmosphere in order to prevent fl'iid pressure from bulfiding up and holding the valve open. Dirmet-acting valves require a large envelope to provide space for the spring and adqmaatment. Alsoi, spring raneSw usually are narrow. As in relief valves, a small pilot section may be added to control the main valve. In a pilot-operatod, pressure-rducing vslve the spool or poppet is balanced pneumatically by downstream pressure at both ends. A light spring holds the valve open. A small pilot rslief valve, usually built into the main valve body, bleeds air to the atmosphere when reduced presure reaches the pilot valve spring setting causing a pressure drnp across the lf a i~v suI "pe.P~.r toward its closed position against the forcle of the light spring. The pilot valve reieves only enough fluld to position the main valve spool or poppet so that flow through the main valve equals the flow reqhah'ments of the reduced-pressure circuit. If no Pow is required in the low-presore circult during a portion of the cycle, the main valve deem. High-pressur air leakinS into the reduced-pressure section of the valve then returns to the atmoosphere through the pilot relief valve. Pilot-operated, pressure-reducing valves generally have a wider ratge of spring adjustumeet 1,1nrity an,~ do rc-.~~ valves, and provide mnor repetitive accuracy. However, contamInatdon cane block flow to the pilot valve, causing the main valve to fail to open properly. 9-3.2.6.4 PreSSwr Regulatoss The pressure compensator bypass flow regulator and moisture separator usually is a part of the air compressr, and controls flow by diverting excess compressor output overboard. In a typical example. flow-pressure drop across ameteoring orifice isused to shift a balanced spool against a control spring. Ths po oeeti sdtomiti osatpa sure drop across the orifice, diverting or bypassing excess supply flow. The pressure drop, which ivdetermined by spool area and spring forme, is relatively low. I-3.2.6.5 Dlreculseal Coutro Valves This term describes all multiple-psassg
valves,
1
S- •three-position.
AMCP70-0 bemcuse their primary function is to control the direct'on of flow from one fluid line to another. Common types include three-way. four-way. divertcr, swquence, and shuttle valves. Actuation may be manual, mechanical, pneumatic, or electrical. These valves are identified by method of actuatio.,, number of parts, number of positions to which the valve can be actuated, type of valving element (spool, slide, poppet, ball. etc.), and type of sealing, A three-way valve (Fig. 9-58) is one with three external port connections and is either two- or threeposition. The usual three-way valve has one common port that can be connected to either one or two alternate ports while closing the nonconnected port. Normally. thems ports are identified as pressure, cylinder, and return (vent). When used to control a singleation c7 cylinder, the cylinder port ;s the common port, and is connected alternately to the pressure port and to the return port. A four-way valve (Fig. 9-59) has four external port connections, which usually are arranged so that there are two simultaneous flow paths through the valve, Four-way valves commonly are used to actuate u-6ao.Lu ;yllindca.in such a'pfi'* ".! valve is connected so that when pressure is applied to one cylisdcr port, the other cylinder port is vented, and vice versa. Four-wa) valves normally are two- or In a three-position, four-way valve, there is a center position in which all ports are vented.
CYLINDER
SI
PRESSURE qI
r
6-
iJ
/j
(A) BLOCKED POSITION
valve, and in this position the leakage clearuince is
CYLINDER
-.
-
e
L
-,
-
•, divertet valve basically is a threc-way valve, with the common port being the pressure port. Flow can be diverted from the pressure port to either of two alternate flow paths. Diverter valves also are called di, version valves. A selector valve functions similarly to a divcrter valve, except that th: common pressure port can be connected to an unlimited number of alternate low paths. A sequence valve is one whose primary function is to direct flow in a predetermined scquence between two or more ports. A shuttle valve is a type of sequence valve that is pressure-actuated in such a manner that whsn a preset system pressure has been reached, the valve automatically actuates, connecting two or more flow path!. A spool valve controls fluid flow by covering and uncovering annular ports with lands on a sliding spool. The number of lands and ports on the spool and valve body determines the porting arrangements that can be achieved, and the geometrical relationship between the lands and norts determines the timing of the valve function. With a sharp-edged land and port. upcrahilun i, i spuuil "'-v- is b-r.u.p. In applications where this arrangercnt would cau.e undesirable pressure surges, the land edges cen be notched, tapered, or chamfered to modify the flow characteris'ics. Spool valvat are c!assifi.d as packed or unpacked, depending upon the sealing characteristics. Packed spool valves use O-ring3 or some other type of seal between the spool and the valve body in order to achieve tight shatoff. Unpacked spool valves possess internal leakage, depending upon the clearance between the spool lands and the valve body. Annular giooves usually are machined on the spool lands to improve lubrication of the valve and to equalize pressure all around the spool in order to prevent binding on one side of the bore. In addition to eliminating binding, the annular groove centers the
-~
PRESSURE
(B) OPEN POSIT:ON rrielement
\Figmsr 9-58. Diren.tial (Cosol Vave--Three-way
minimized. A uniqiie feature of the spool-type, multiple-passage valve is that the end of the spool car. be used as the actuator piston to position the valve. Poppet-type, mpultiple-passgts valves use two or more flat, conical, or . •herical seats on a translating poppet. These valves lend themselves to three- or four-way operation with a variety of seating and sealing arrangements. For a solenoid-actuated, threeway, poppet-type valve in the de-energized position, pressure is applied to the cylinder. "OThe sliding plate valve consists of three main element3, a slide and two plate enclosures. The slide
is san~lwiched between the two plates, and ontains co-ed holes and passages that mate with 9-39
AMCP 7W6202 CYLINORJ 1
I
-
CYUNOE3 2
IRE
-•
TUN
CvLIi PFFs9
RE.T
Cv.YL 2
(A) POSITION I
ports in the plates. Many porting arrangements are available, and thiee- and four-way mulitpl.-pssuge valves can be &chived easily. One advantage of a sliding plAte valve is that it can be reworked and lapped to compensate for wear. A ball valve cat, be adapted readily to operate as a multiple-passage valvc by the aldition of outlets on ;he body and additional porting in the bell valving element. With three outlet connections on the body, the ball valve can be made into a variety of three-way valves, depending upon the porting utilized in the ball. A rotary slide valve is usrd more commonly as a multiple-passage valve than as a two-way shutoff valve. Rotary slide, multipl-pasage valves consist, aesntially, of two parts, i.e., the body and a rotating plate (Fig. 9-60). The body contains either three or four outlets to provide a three- or four-way valve configuration, and the rotary plate contains various porting arrangements so as to achieve a multiplicity
of three- or four-way valve types. A valve actuator is a power unit that proviies a mechanical operating force for positioning a valving element. The actuator may be eith-.r direct-acting cr
-
piloted. The direct-acting valve actuators include S
Rinvolving
largcr power source, is common with such pistoncylinder actuators as the electro-pneumatic., pnoumatic-hydraulic, or pneumatic-pneumatic w.•mbinstions. Most aerotracc valves are powered by rtmotc-
LI P•F.
W
o
vo torque motors, each used indepcadently. Piloting, the use of a small power input to control a
rCYL 2
control, or automatic, actuators
Linear actuators
used in aerospace valves include solenoids, pisto.cylinders, bellows, and diaphragms. Rotary actu-
(B) BLOCKED POSITION
ators also can be used to impart linear motion, for example, through a rack and pinion. Linear actuators aiso can be used to impart rotary motion by driving an internally threaded valve: stern attachment "with a threaded rotating shaft. The actuator may be
either an integral part of the valve, or a separate PRESSURE
Ii
R
/EýLýC P
I CYL
REthrough
PRESS
RE.T
device linked mechanically to the valve. Solenoids, diaphragms, bellows, piston-cylinders, and servo torque motors are usually in, :ral parts of the valve. Electric motors commonly are linked to the valve a gear train, s*r•w drive, etc. Valve actuator positioning requirements can be divided into two gh-oupb:
CYL 2
(C) POSITION 2
Figure 9-59. Directional Control Vahe--Foar-way
9-40
I. Two-position, or nonmodulating, actuators for "on-ofr*control valves and shutoff valves
2. Modulating actuators for positioning control valves; pneumatic and hydraulic piston-cylinder actuators commonly are used with valve positioners, and provide a feedback loop between the control
I
)
AMCP 7M6-_202
/
CYLINDER
2 IN
CYLINDER
CYLINDER
L CYLINDER
.S PRESSURE
PReL [, IR•2E
PRESSURE I
riR
SSURE S.--:"-PRE
SLIPOSITNION
I
PRPRESSURE ImLSSrL
Z\
-
CYLINDER
"2 ICYLINDER SLDE ROTATNG
J CYLINDER
BODY
PRESSURE
PRESSURE
2
i
4
1
'
I'-"J
/2
-
.nt1t BLOCaLLI
POSITION
BLOCKED POQi TION
PRESSURE PRESSURE
CYLINDER
CYLINDER
7
-
~2\
CYLINDER
2
-
C C-LIL REssu
IPOSITION
2
IPOSITION 5.
PRESSURE
VALVE SHOWN IN BLOCKED POSITION F11m 9,40. Dirweeual C"Mle Vahte--Rotary-Four-wsy
V
.
"•
Oaln and the stroke of the actc•nor. The design or selection of a valve actuator is. at best, a trade-off among several interrelated factors Typical mechanical factors that must be considered indude length of stroke, locking requirements, speed. limitation of envelope size and weight, and maln,tude of required force. If a long stroke is needed, use of solenoid is eliminated automatica!ly. If fast .response times are desired, actuation by electric
motor, solenoid, or explosive charge should hle considered. If high forces must be overcome. either hydraulic or pneumatic pressure must be used. ".2.7 PreNMre Gages Pressure gages are used in fluid-power equipment to provide: 1. An indication of operating pressure, especially where this pressume must be selected by the optrator Q-41
Q_
AMCP 706-202 "
r
:•
,•
-g Stirs
S~4.
2. An indication (alarm) of abnormal pressure within whts the system..Air Pressure gi.,es also are used to provide data in ,development of fluid-ower equipment. Pressure, or pressure chunge, within a system must be correct if pneumatically powered or controlled equipment is to operate properly. TI.e proper gage itdicates this pressure and helps to prevent malfunctions. Gages also can be calibrated in valu.-s proportional to pressure, ~~~~~pnieumnatic cylinder. such as total force exr-rted by a =ourdon-tube indicating dial gages (Fig. 9-61) are uswzd to measure pressure from 0.5 psi vacuum to 150,000 psi. Primary advantages are accuracy, ruggedness, reliability, simplicity, and low cost. Other methods of measuring pressure include electronic :.• devces ev-es based ase upn eadags or upon srai strain gogo gage read'.igs. or, inthe in the case of pulsating pressure, piezoelectric crystals. Such units, relatively, are costly and comple'x. At thc low •end of the pressure-measurement spectru are bellows and diaphragm-type devices, which are used to measure relatively low pressures. Components of Ji Bourdon-tube gages are similar. 1-Howevei, many styies and materiais should he con"sidered ii. selecting a gage. The following factors, listed in the normal order of consideration, influence hi selection: .!. Measured medium, including pressure range -and fluctuation 2. Etivironmental conditions, such as temperature and vibratio i 3. Wear conditions ,:aused by pulsation and vi.bration Connection of gage to measured medium 4. Counting oethoe 546. Muizehng method "6. Size and weight 7. Accuracy.
a
TUBE
Air Storage Bttles storage bottles or vessels are often of conven tional shapes, such as cylinders or spheres. On the other hand, limited space may require a conical, oblate spheroid, t,-roida!, or pear shape. Drawn cylinoers up to 9 in. in diameter and up to 50 in. lone are available. Welded containers can be made much largcr. Capacities may range from 3.0 to 3500 in.' Cylinder bottoms may be concave, bump-shaopd, spherical, or elliptical. Ports located wherever peiao litcl ot can a beelctdweee they are required. Either external or internal threading can be supplied. Pressure vessels with pressure ratings varying from L,few hundred psi to 25,000 psi are available. Pressure vessels can withstand high ambient temperatures in the range of 2750-6W00F. In addition, they can m.t vibration, shock, and other extreme requirements of
v odn shcknthen retme reqdirhoice of rmodern helicopter environments. A varied.hieo constructions is possible. Am'•n5 the metals that have been used successfully are alumiatum, low-alloy steels, high-strength steels, and a number of exotic metals. For highly predictable, multicycle performance, one of the chromium steel alloys is recoi-
Seamless cylinders are deep drawn, and wall thickness tolerances can be held preisely. An excellent surface is obtained. The dome and Peck are hot spun on the open end. Heat treatment produces thz mc.allurgical quality nr eded for best performance. Seamless drawn cylinders are relatively inexpensive io produce. Welding often is used to fabricate the larger sizes of cylinders. Certain smaller cylinders for aerospace applications, where minimum weight is a more im.ortant consideration than is price, also are wclded. Containers with a wide r!-nge of wall thicknesses, diameters, and alloys can be welded, as can vessels of exotic high-strength metals. Fiberglas vessels usually are lighter than their allDIAL /POINTER metal counterparts. Advanced technology has made them highly reliable. Gilrne-finished glass, bonded with epoxy resin and protected ,gainst moisture. BOURDONpenetration by ar. external coating, can be used to fabricate pressure vessels. Internal rubber linings
/
ELINKAGE
MECHANISM SCASE
PRESSUPIE TAP Flgir'-. 9-61. Pressure G.ge-RBourdon-tube Type 9-42
9-3.2.8
ffectivcly retain the air in instances where some permeability Composite can be tolerated.
cylinders offer extremely low weight
without sacrificing reliability. In this type of construction, a cylindrical metal shell having hemispherical ends is wrapped with circumferential weldings - usually of bonded Fiberglas. Hoop loads are shared between metal shell and windings while longitudinal loads are carried by the metal alone. Wire winding may be applied in accordance with
(
'I)
AMGP 7W0~2O
Miltasy Spacirications in order to keep vessels from shatterift under gunfire. Wire winding contributes added 3traigtb, but usually is not regarded as a light-
Paua*isa
ess
hc
u
hre
niaterial, shape of the vessel, surface conditions, joint design, heat treatm~ent. and environmental conditions during urse must be contidered. The ratio of test pressure to msvice pressure is usually 1.67:1. Howb ~r, the ratio is governed by individual speciricatioma and, in some cames may range from 1.5:1 to 2.0: 1. The ratio of minimum burst pressure to service pressure depends tipon the design stress levels, but is commonly 2.22:1. Ratios of 2.0,1 to 4.0:1 often are indicated by specific cycle,-life requiirments and other 9-3.9 empeues(C) ssysre. Subsystem components are discussed in p~s 9-2 3.2.9.1 through
U4JM*'
(A)SINGLE ACTIKI~
(B) SPRING RETURN
DOUBLE ACTING ,k
I...A#aO~y~ -%6aa~
9-3.2.9.1 Aecuaters An actuator, as used in helicopter applications, is a power unit that produces a force or torque for positioning loads. Notrmally, pneumatic actuators are of the linear-motion type, anti are designed to individual specifications. Among the types manufactured are specialized actuators of the piston type (with built-in dampers). used for the retraction of ~lianding; gear; landing gear up-lock actuators; highA temperatuare piston units for both high and low pressures; cargo ar.d passenger door actuators; storeejection actuators; screw-jack actuators for high-ternperature appficadior. 1 and air motor and screw-jack actuator assemble.. that form a part of such systems asnosewheel swrving. Advjantages of pneumatic acJkNtuators include speed of operation, simplified power requirements, and ability to withstand ambient toniperatiares to 500"F. The inherent limitatjons of pneumatic actuators result primarily from th-e elastic properties of the compressed air woa king fluid. Fig. 9-62 illustrastes typical lintar pneumatic actuators. one direction only, and can be cithee the out-stroke or in-stroke. The return stroke is accomplished by ~im~som esernal means; a double-acting cylinder zan be usdfor this purpose by connecting the ac, iating fluid line to only onz port through a threeý-way valve, 'Nleaving the other port open. Special single-acting cylinders are desirnee with piston-scalinS devices; _
that seal in one direction only. These cylinders have a po.-t hole in one head and a bleeder hole in the opposite head,. The spring-return actuator is a single-acting cylinder, with the re'uarn stroke effected by a spring. The length of the cylinder in the retracted position is nt least twice the actual stroke length because of the spring length. The initial spring force, as well as the increase in spring force due to spring rate during compression, depends upon the amount of oc required by the spring-actuated stroke. In the double-acting actuator, the cylinder has a power stroke in both directions. The actuating fuid line is connected to both heads of the cylinder, usually through a four-way valve. Most standard catalog cylinders are double-acting. The sealing aevices, also operate in both directions. Rotary actuators rotate an output shaft through &A fixed sate to produce oscillating power, converting fluid-energy input to mechanical output. Thy r ducing higb instantaneous torque in either direction, and requiririg only limited space a~nd simple mountings. Rotary actuators consist of a chamber or cbaaabers for containing the working fluid, and a movable surface against which the fluid acts. The movable surface is connected to an output shaft to produce the output motion. 9-43
Z
~-
-N
tl
The basic types of rotairy reciprocatins actuatoms cmaw aM Tog pistor.-rack u-tuator may have a are vane and piston. Basically, the vane actuator concylinder witL4 two pistons, each integral with a rack, sists of a cylindrical chamber, a stationary barritr, a or two cylinlers with four pistons. In each cade, the cenitral shaft with a fixed vane, and end caps throug~h racks engage a pinion in the -rntcr of the cylinder and, as the pistons move the racks, the pinion is rowhich the s~haft projects and which support the shafttated. Equal torque thus isproduced in each direction and-vane assembly. Fluid energy on one side of the vane produces an unbalancW force on the shaft. The of rotation. fitom usually ame sealed by standard 0ring seals. Units are aivailaboe with higb sorjue shaft extension can protnado from either or both ends, and has key-ways, splines or squared ends ior rat~igs and for preaourew to 300C psi. mechanical connection of the hlad. Vane actuators usually have one or two vanes. but may have three or ,0*, aim* more. The amc of rotation for siraglc.vant units is emaibrevlesrodeamns f about 2140 dong; o obevn ntaot1og supplying operating pressure to lacIhpter wheel Maxium aris wth rc he izeandconi utio of brakes. The pressure applied to the pedas (or hand the unit. Besides the torque, there is a radial force Oisle rdpniguonvleesn)divr& the vane shaft, and this side load tends to deflect the regulatod, proportionate pressure to the valve outlet shaft. Thus, the output shaft must be large enough to port thus alflowing sir to flow to the brakes. withstand mrximurn torque and side load withoi t Releasing of the brake pedal (or handle) vents the excessive deflection, and bearing areas must be large pressurized air on the brakt*. Pedal or handle travel enouh t suportthes lods wth inimm war is proportional to brake valve outlet pressure. i.e., to and friction. EMTciency of single-vane units varies the pressure applied to the wheal brake. The design of "Tom '1D to 95%, depending; upon such items as brake valves should feature minimumw hysteresis charbearings and bearing length-to-diameter ratio.uitc.Frnresdeeaeepoeaiflainhia;R-4
A.dard
Mines.ai:.. ripA-
exhaust
vav
are located between
th L-.k
valVCS
posing stationary barriers. Fluid enters one compaitand the brakes. nient from an external port and flows through internal passages to the opposite cornpihtwent. A force'i exerted on eacb vane, and the force that tends to disA fuse Paisusdtopteca F pneumtcsseshl place the shaft is balanced transversely. Because pure Afs sue opoe~apemtcsse hl peitigflffcetoraonfcopntzn torque isthe only load on the shaft, cffidency of the the remainder of tht system. Eisaitially, thc fuse 'double-vane unit is high; torque output isdouble~ that operates on a rate of flow that is sensed by a pressure of a single-vane unit of comparable dimensions. rptruhtefs tsf necoetefs Vanle-type actuators are available in a variety of Stansizes and mountings. with torque outputs eesatmtclya eo;stepesr ifr ranging from 3 lb-in, at 50 psi to more than 7,000~eta srmvdb enigteusra on Piston actuators are available in several types. In a helical spline actuator, fluid is applied to one side of thec piston which 4 kept from rotating by guide rods. The actuator can be stopped at any point in its stroke. The: helix angle on the shaft and piston isselflocking, preventing rotation of the actuator under external torque loads. Sealing is by ring seals arounO the piston, the fluid rods, and the helical screw. Standard units are available for a wvide range of torque outputs and pressures. The arc of rotation can be larger than 360 deg. Adjustable cushions art available to reduce shock at each end of the stroke. A variation of this type of a-,tuator uses two pistons in the same cylinder. Tle shaft has a rightha id helix on half of its length and a let'. hand helix on the other half. Fluid is introduccd into the area between the pistons and causes :inear movement of the pistons, thereby imparting rotary motion to the 9.44
9-3.2.9.4 Qulk-1somets Quick-disconnects provide easy, instant coupling and uncoupling of pneumatic systems and system components without loss or supply pressure. Their use greatly facilitates aircraft overhauls and servica replacements. In order to 'meet varying requirements, two types of quick-disconnects are available: a lever type and a rotatin~g type. PNEUMATIC SYSTEM INSTALLAYION AND QUALIFICATION Pneumatic systems are classified into types and classes as follows: 1. Types: at.Type A. Airborne compressor-charged system, in whicn system air pressure is maintained by a compressor mounted in the helicopter 9-3.3
(
) h. Type B. Ground-cljarged syrtem, in which systakn air pressure is obtained from ground-servicing equipment 2. Claism: a. Class 1. Supply system is -;harged to a pressure of I 500 tai Class 2. Supply systnim is charged to a pressure of 30W psi c. Class 3. Supply system is charged to a pressure o1F 50)00 psi. The qua~Lfication testing required for Type A and components is similar. The tests indlude those for exanmination of product, proof and burst pressure, leakage, flow and pressure drop, extreme temiperatrlife cycle, vibration, humidity, fungus, sand and for. and dielectric strength. General requianets orpneumatic-system component ttsting are given in MIL-P-8564. The conditions specified should include the test media, temperature-s, and il1tration. System installation testinf. requircivenrts are listed in MIL-T-5522 and AMCP 706-203.
-*b.
1'
AMCP 7OC-202
-,B
V~ \_A~dustments
exist and the altiniewc setting is 29.92 in. Hg, the altimeter will read the correct fielt4 elevation when the helicopter is on the ground. The altimeter in its simplest form is shown in Fig. 9-63, aid consists of an evacuated diaphragm or capsule mounted in an airtight case or static-pressure chamber. The diaphragm responds to changes in pressure by expai ading and contracting. and the movement of the disphrAgm is trcnsmitted to a main pinion assembly. The dial is calibrated to mead pressure altitude. The static pressure hneasured P. at the static source of the altimeter may differ slightly from the true atmospheric pressure p. For any P.. the altimeiter, when corrected for instrument error, will indicate the measure-d pressure altitued corrected for instrument error H.. The instrumrent error is a~l error built into !he altimeter, consisting of such things as scale error. The quantity (Pm,- p)l is called the static pressur-. error or position error and is determined through flight tests. The value that is added to R. to determine true prtr~ure altitude H is termcd the attimeter position error correction 6 H
9-3.4 PITOT-STATIC SUBSYSTE:M DESIGN ftreim~p aoitjitmrdu-,ng
rptp nf climh are-
basic parameters in the performance of all helicop-
Xters.
j
Instruments used to measure these quantities are the airspeed indicator, altimeter, and rate-of-climb indicator. The pressure inputs to the airspeed indicator are obtained from the pitot tube, which niaue~total press ire, and fromn the static pre- ure source. The latter also provides pressure foi- the altime~tCT and rate of climb indicator. '-he pressure sources on the outside of the fuselIage are connected directly to the instruments in the cabin area by means of leak-tight tubing. Design and con-
A
absolute-presure gage, called an altimeter. )sitive wcaled so that a pressure decrease indicates an alti-
t ide increase in accordance with the U!: Standat d Atmosphere (Ref. 5). If mnandard atmosphere conditions
..
Li
-
A,
where - true pressure altitude, ft Hf - measured pressure altitude, ft H, ZýH- - altimetcr jxosition crroi correction, ft Minimum performance standards for a pressureactuated, sensitive altimetei r ie nFAIO Cl Ob. There aiso aire a number of Military Specifications available that cover specific altimeters currently used by the military.
MIL-P-5518 and MIL-P-8564. No valves or severe restrictions art permnitted. Drain fittings shallI/be pro--vided, as necessary, at low points in the sys .emn in order to permit removal of condensed moisture. Military Soecifications governing pitot and static systems on all airzaft and missiles arc MIL-P-26292 and MIL-I-61 IS.Thesw specifications are written primarily for conventional aircraft, but are applicabie tocompound helicopters having alternate means of producing horizontal thrust and, therefore, increascd forward airspeed capability. 9-3.4.1 Altiumetlers Most altitude meaurements arc made with a scn--
,.,
--
I
STAT IC 0 '.-
ESt
P
,
7
Figure 9-63. Altimeter Scematic 9-45
AM!C
7OW2O
The design and installation of the altimeter system shdal te such that the error in indfated pressure altitude at sea level in standard atmosphere, excluding instrument calibration error, does not result in a reading mere than 30 ft high nor more than 30 ft low in the level flighi speed range from 0 mph to 0.9 times the maximum speed obtainable in level flight with rated rmm and power.
*.•
9-3.4.2 Rate-of-edlml ladlcator The rate-of-climb indicator uses the same static pressure source as the altimeter. A calibr'sted flow restriction is placed in the unit to restrict the passage of air into the instrument case when the measured static pressure changes. The time lag associated with this pressure change is used to obtain a pressure differential on two sides of a diaphragm. This pressure differential is displayed mechanically as change in "pressure altitude with time dIIdi at standard sea level conditions. The re:ationship is in accordance with the hydrostatic equation for small differences. AP - pa,(A ),
lb/ft"
t9 -2 )
where
ac.leratinr edue to gravity. ft/sec2 - height ditferencc, ft - pressure difference, lb/ft2 - density at standard sea level conditions, slug/Its The slow response time of those basic mechanisms is now generally corrected by the incorporatior of accelerometers which provide an artificial boost of air for an instantaneous needle movement. AH AP PSI
9-3.4.3 Airspeed Iickators True airspeed V is the velocity of the helicopter wits' rcapc ,.t#-.,, a.r through which ist flvin. it i% difficult to measure true airspeed directly. Instead,
used in establishing the table. Airspeed indicators are calibrated according to tese relationships. In operation, the airspeed indicator is similar to the altimeter. However, instead of being evacuated, the inside of the capsude is connected to a total-presure source and the case to the static-pteuaure source. The instrument then senses the differencs between m%asured pitot (totul) pressure P,,, within The capsule and measured static pressure P, outside as shown in Fig. 9-64. The pressure differential q, (Eq. 9-4), as measured by the airspeed indicator when correted for instrument errot, will correspond to measured calibrated airspeed Vo,.. qo. - Pt. - P,'. Ps (9-4) where P,, - measured static pressure, psi measured pitot premure, psi measured impact presu'ire. psi In general, this will vary from the correct calibrated airspeed because of pitot and static pressure errors, and an airspeed position error correction A V, must P,.
be added to V,. to obtain V,. A V, - V, - V,,, kt (9-5) whee, VP - cal.brated airspced, kt ., - measured calibrated airspeed, kt A V, - airspeed position error correction. kt To determine its significance, this correction is evaluated by flight testing of the entire airspeed system. Specifications have been written to establish the acceptable instrument errors and general coastruction ot airspeed indicators. Table 9-2, AMCP 706203, specifies requirements for airspeed indicatots. Because of the ccmplex flow patterns around the helicopter in flight, the airspeed system must be flight " -
--
-
calibrated airspeed V, is measured. Calibrated airspeed is determined froin the difference be'ween total pressure and static pressure using Bernoulli's cornpressible equation for frictionless adiabatic (isen-
-
-
-
tropic) flow. Calibrated airspeed is the adopted standard reading of an airspeed indicator, and is the same as true airspeed under standard sea level conditions. The difference between total pressure P, and static pressure P is called true impact pressure q. q,
P-
, psi
STATIC PRESSURE
(P.)
(9-3)
where
."
P
- static pressure, psi
P,
- total pressure, psi
- true impact pressure, psi cc A tabulation of q, as r function of Vc is given ir. Ref. 6 t oth. with a complete derivation of tIar. equations 9-46
•
...
TOTL R SUR TOTAL PR(ISSuE
Figre 944. Alrope
lUllater S*bsalek
..
9
~AMCP 7W6202
tesmed An order to evaluate performance. The flight teat for airspeed calibrrtion is described in par. 9-5.2, AMCP 76-203. This night testing must be acc&'mplishe over the full range of flight capabilities of the helicopter, including climbout, cruise over the entire speed range, aitorotation, and rideslip. Flight test teoniquas, established primarily for conventional airak are de-cribed in Refs. 7, 8, and 9. Some specific flight test instrumentation for V/STOL aircraft is described in Ref. 10. 9-,4.4 Tmdal-prssmre Sowees to determine or pitot-static total P~tot or pitot presure, tubes whichare isused defined as the prisure of the air when it is brought to rest isentropically. True total pressure is measured by a sharp-lipped pitot opening that is faced directly into the airflow. Pitot pressure errors can develop when the airflow impinges the tube at an angle of attack or incidence. A wind tunnel investigation of a number of total-pressure tubes at high angles of attack is desc•i•bd in Ref. II. Results indicate that a sharplipped opening with shallow or no internal and external tapers has the least sensitivity to angle of attack,
4
The nitnt nni_. _nihi. €
)
vn
-
in-
flotw pulsations. Locations on the forward top of the canopy or above the rotor on a stationary rotor hub have been found satisfactory. Rotor downwash can cause pressure pulsation in a pitot tube. Ref. 12 evaluates the effect of these pulsptions on total pressure. Proper design of the pitot tube and connecting lines for pressure-lag response, together with a favorable mounting location out of evre dotrnwash, can eliminate the pulsation probbIn. Flight test investigation of prosure-lag problems is docum•,fted in Ref. 7. Although ov. pilot tube is sufficient for obtaining total pressure, •t is recommended that, if there are two sets of instruments (i.e., two airspxd indicators). separate pitot iources be used for each indicator. This redundancy permits detection of a faulty pitot pressure reading caused by a plugged or iced-ovc. pitot opeaing or a pressure leak in the system. s oureSo 9-3.4.5 Satk Presea A number of helicopters are designed with flush static prewsure vents located on the fuselage. These static vents or ports shali be located such that vehicle speed, the opening or closing of windows, airflow ,v'-" ca.
,,,
w- tiv w viir tr-i-n ma-rii
to place a -' ation. Therefore, it usually is desirab!e water-collection chamber and drain hole in the pitot
altimeter, not affect their accuracy seriously. Each airspeed indicator, and rate-of-climb indicator shall
tube. The small amount of airflow that passeb through the pitot tube and out the drain hole also
be connected into the system in an airtight manner, except for the static vents.
causes a pitot pressure error, and should be investi-
Two static vents normally comprise the static
gated for each design. Helicopters can and do operate in atmospheric conditions that are conducive to producing ice, e.g.,
pressure system, They are located symmetrically on the right and left sides of the fuselage, and are interconnected to the flight instruments by a single tube
from supercooled water droplets in the air or from freezing rain. The tip of the pitot opening, because it is in a stagnant-flow region, is susceptible to blocking ..iy -- d can c .mecompt.ly plug' ,-and and operative even before an appreciable accumulation of ice has developed on the rotors and other parts of the vehicle. It. therefore, is important to use electricially heated pitot tubes that are capable of deicing and anti-icing under the most severe atmospheric condittons that are likely to occur in flight. Pitot tube. normally are pisced in the forward area of the heli opter, outside" of the airflow boundary layer. Loca ions high on the fuselage are desirable in order to avoid ground damage. Tht pitot tube shall be pointed into the nominal flow direction, and should not be placed behind any protrusion thlit could cause flow separation ahead of the tube. It also should not be located aft of windows or openings that couid exhaust airflow into the pitot opening. It shall "belocated where the dowawash from the rotors does no: cause large-flow angles of attack and excessive
containing a T-finting located halfway between the two vents. This right- and left-hand installation is used in order to reduce sideslip errors in the system, also provides a partiai redundancy if one of the vents becomes plugged or damaged, or if a leak develops. For helicopters with advanced speed and altitude capabilities, a leak to a pressurized cabin area could cause a serious static pressure error. In this case, two completely separate static systems with separ,•te right and left vents should be used if there are two sets of flight instruments. Common locations for static vcnts are on the sides of the fuselage aft of ahe cabin or on the tail boom. A location rhall be selected that is n.ot sensitive to rotor downwash or forward speed. and that is away from windows, doors, or air vents that could produce a variable airflow geometry in the: vicinity of the vents. A flight test program is necessary in order to determine the pressure influence of all vat iables, in addition to the calibration under c•onditions of climbout, cruise over the speed range, and autorotation. 9.47
\
..
,
AMP7W6202 Static ventsAlsllbe located so that no moisture can enter t~,e openings under any service condit'ons. The Static vent plate shauld be heated if there is a pro~.iability that ice could aeal over the static vent. Other pertinent design information for flush static vents can be obtained #'vrnnMIL-1-6115S and MIL-P-26292. 9-3A PIWftdcTabu1957. On recent helicopters, the pitot and static pr'essure sources have boan combined into a pitot-static zube. This tube ciui be straight for boom mounting, or Lshiaped f ýr mounting directly to the fuselage. Acceptible mounting locations are on a short boom ahead of the nose of the helicopter on a stationary hub onl the top of the rotor, and on the top of the cabin toward th: front, yet near the rotor hub. L jcations on the forward sides of the fuselage also are acceptable. However, lo% locations on the bottom of the fuselage usually are susceptible to ground damage. If the tube is located at the rear of the fuselage, rightand left-mounted units should b,. used in order to reduce the influence of sideslip on the static pressure measurement. For redundancy, two pitot-static tubes muizmummonded for an heiicopters having ýwo sets of flight instruments. Pilot-static tubes also rnan be designed with two sets of static ports if additional static sources are required. Minimum requirements for pitot-static tubes are specified in MlL-P-31 136. It is recommended that all pitot-static tubes bc capable of completely deicing and anti-icing under the most severe envitonmental conditions likely to be encountered in flight. Placement of the static ports on a pitat-static tube also assures adequ~ate deicing of the static vents. Pitot-static tubes offer the possibility of aerodynainic compensation for static-pressure errors. This is accmiplished by selectively designing the shape or the tube and the location of the static pressure ports. A general description of aerodynamic compensation isgiven in Ref. 13. The concept isused extensively for conventional aircwafP, both commercial and military. MIL-P-83207 applies tor straight-boom-mounted, aerodynamically compensated pilct-static tubes, and MIL.P-83206, covers L-shaped, compensated tubes, Placement of the static ponts or the pitot-static tube and away from the fuaclRSe skin also can reduce errors caused by iocal skin irregularities in the vicinity of flush static vents. An extensive investigation of surface irregularities and nonreproducibility has been peaformned for conventional aircraft (Ref. 9). REFERENCES 1. Blackburn, Reethoff and Shearer, Eds., Fluid Power Control, The M. 1. T. Press. 9-48
2. B. W. Anderson. 77e Analysis twSs Deuignv of Pnewmatic Sysiut ,tJohn Wiley and Son*, Inc., NY, 1967. 3. Michlne Designv. Fluid Power Reierence Isaue. Sept. 22, 1966. 4. FRow of Fluids Throulh l'alw~s, Fittings and Pipe. Engineer Division of Ctane Co., Chicago, IL, 5. U. S. Standard Atnsophere. Prepared Under Sponsorship of National Aeronautics and Space Administration, United States Air Force, United States Weather Bureau, Washington, D-C, December 1962. 6. Sadie P. Livingston and William Gracey. Tables ofAirspeed, Altitude, and Mach Number Basedon Latest International I'aluesfor At:mospheric Prop-
ere and Physical Consta.,ts. NASA Technical Note D-822, August 1961. 7. Russel M. Herrington, et al., fl7ight Test Engineeting Handbook, AF Technical Report No. 6273, Air Force Flight Test Center, Edwards Air Force Base, CA (AD-636 392), January 1966. 8. Richard V. 1>,Lco. Peter J. Cannon, and Floyd W. Hagen., Evuaiaadoa of New Aes'hod~i fo Flighi Calibration of Aircraft Instniment System,. Part IPAnalysis of Altimeter. Airspeed, and Free-Air-
Temperature Systems. Wright Air Development Center Technical Report 59-295 (Rosemount Engineering Company Report 6591). June 1959. 9. Richb.id V. DeLeo. Floyd W. Hagen, Robert R. Kooiman, and Donald 1. Thompson, Evaluation of Factors Affecting the Calibration Accuracy of
Aircraft Static Pressure Systems. USAF Technical Report SEG-rR-65-35, Air Foroe Systems Command, VWPAFD. OH-, (Rosemount Engineering Company Report 66227). June 1962. 10. Floyd W. Hagen and Richard V. DeLco. Flight Calib atlon of Aircrat Static Pressure Systems, Federal Aviation Agency, SRDS Report No. RD-.4i-3. ("losemount Engineering Company Repart 76431) February 1966. 1I. William Gracey. William Letko. and Walter R. Russell, Wind-Tunnel Investlgauioai of a Number of Total-Pressure Tubes at High A ngles of A ttack.
Subsonic Speeds. National Advisory Committee for Aeronautics Technical Note 233 1.April 195 1. 12. R. C. Johnson, Averaging of Periodic Pressure Pulsations by a Total-Pressure Probe, National Advisory Committee for Aeronautics Technical Note 3568. October 1955. 13. Richard V. DeLeo and Floyd W. Hagen, The Use of Aerodynamically Compensated Pitot-Spatic
Tubes on Aircraft. Rcsamodnt Engineering Coinpany Report 2686C. Minneapolis. Mi.
(-
AMCP 7*-20
CHAPTER 10
INSTRUMENTATION SUBSYSTEM DESIGN The lighting concept shall consider the total cockINTR~ODUCT'ION pit instrumentation rcquirernieits rather than inTh'ni chapter discusses the instrumentation neeneeds. The necessity to consider all requirc. dividual "0*WY lIn the helicoptor cockpit to permit assigned misMents cannot be overemphasized. Also, becaus the eeoeruuly rcrscoki n am sions to be performed. Because missions must be coniductd at night. or urnde IFRK conditions, the require strumrntatio-i from many different sources, he mtst ments for lighting of the instrument and control pan recognize his role as the cockpit lighting integrator els an indkuded. early in the detAil design process or the resulting The instruments requieed in order to provide the cokpit light',.g will s~affer from problems of varying pilot with the information niecessury to conduct as-colors, imbalance, and unevenness eý,cn though all signied mission&normally should be grouped by runcvendors art designing to the same requirements. Con. tional categories. Indluded are flight. navigation, helitrol of cockpit glare and reflections also must be adcopter subsystem, and weapon system instruments. drsebytearinevlp. lnstaliation roquirmennts for instrumentation ame diedresse byrthe primarym developer. lgtig h Ecp o h rnayisrmn ihig h cussed! separately. Decattse of the htgh vibration encockpit lighting for all helicopters shall be designed in acodnewtthaplabeposinofMLvironment in which helicopter inst:uments operate, they shall be designed and qualified to survive curve 6503. Integrally lighted instruments having a white M (SG's) of method 514 of MIL-STD-8 lOB. lighting system shall be dcsigned according to MILThe definition of the total helicopter instrumenL-27160. MIL-L-25467 shall be used if operutor retation package sheil include the arrangement of all dispaysi nd ~ cntrls ~ AifC~ ~quircmenis dicmnic inc use oi red lighting. An dislay. ad cramh, i, me cbeI located sinfiasWiches, radio controls, auxiliary controls, and citinstrnuments and displys shall bloaeinccuit breaker panels shall be illuminated by plasticcordance with MIL-STD-250, unless unique mission plate, edge-lighted panels, as specified in MIL-P-?788 requremets nd sppot~nghuma fatorsenand MIL-L4 1774 for red lighting or MIL-P-& 3333 gineerirX (H FE) analysis dictate otherwise. Displays frwielgtn.Tescnayisrmn ae ofa1nfl syteTheprovidearminimrumen frwelighting. for other crew stations shall be arranged to provide lihtndleysofetherhl preorvwite illmintionm on 1 fthugood control/display compatibifity, along with mininthsr inswhtru en iluianel. m ren facdesof ether for mumi workload and minimum' oppoilunrtiy human error. The most satisfactory method of assuring that these goals are achieved is by a systematic 10-2i LIGHTING INTENSITY CONTROL HFE analysis which should be accomplished as early ! isdis~used n mre Unless Gtherwisc specified by the procuring acin he esin osabl. oaseas revihew adei paevauo asrosbe: dtiscu issedinCatrl 5, tiuul0aibl ihigitnst otosa tivity, the instruments shall be grouped on condetai in par. 13-3, AMCP 706-201. Formal mock-up
10-1
mock-up is not limited to the evaluation of the in-. stnzniet subsystem.
REQUIREMENTS
\
1)
WIH~ GENERAL The lighting concepts in use today vary from red to white in the color spectrum, and from direct lighting to diffused indirect lighting techniques. A number of solid-state light-emitting instruments recently have been developed, and these may be appropriate for use in helicopter cockpits. Such applications have used electroluminescent (EL) lighting, liquid crystals, and light-emittingl diodes (LED).
I. Side-by-side: mana. Pilot's basic flight and navigation displays on
plays on main instrument panel
c. Propulsion and other subsystem displays on main instrument pancl d. Center console e. Overhead console f. Secondary panel lighting 2. Tandem: a. Pilot's basic flight and navigation displays on main instrument panel
b. Pilot's propulsion and other subsystem dis. plays on main instrument panel
c. Pilex's side consoles
d. Pik-t's secondary panel lighting a. Co-pilot/gunner's main intrument ptmel
f. Co-pilot gunner's side consoles g. Co-0ilot/1unner'a seondary panel lighting, Lighting rheostats sW1. be canxble of continuous adjustment from FULL "ON" to FULL "OFF" to provide the low settings requit&l for use with light amplification dvioss such as night vision goggles. A single switch to control all cockpit lighting AII be provided if night vision g0$&le will be used extensively. 1."2.3 LOW INTENSITY RlADAJILITY Particular attention must be directed toward optimizing the primary instrument lighting and edge-lit panels for readability at low intensity settings. All pointer, scale marking, and nomenclature sui be readable at 30% rated voltage when viewed from a distano of 32 in. under fully dark adapted conditions. This can best be accomplished by mainStaingq utnifcrmity in sale design. and letter size and significantly in the may differ font. Two instuments ., ..... %:., . .. I.:.... =d th,. , meat with the grater total marking area wid appear brighter. Caeful attention to balancing the ar between instruments will reduce the difference. In some cam a resistor may be added in the lighting circult of the instrument with greater apparent brightnms if matching the arem of the dial face markings is not pinatical. Additional design gwuidaiue in instrumeint lighti2 design may be found in Ref. 4 and 5. 10-2.
WARMI G, CAUTIOi, AND ADVISORY
All waraing, caution, and advisory signals "! • •Ml ~ diqilayed, t.As~vi~qali designed, sad operated inaalnauaebly accrdanoe with
dition. The use f avoioawarning systan (VWS)am
should be considered. VWS is advantasous perticularly during mission phaose when the crew's task
loading is high and their attention is directed ouluide the cockpit. Under such conditios, a light mIga fIraquently may go undetected for long perlods of time. In addition when a problem is deteed via a ir" caution or wgarning li'ght, th speifi Maim, or warning light then must be located and mrad, d otrective action initiated or deferred. dspending the cuiticality of the problem. With VWS. thi cow is made aware immiwiwtely of the exact atme of tdo proolem and can decide whether to niltiate or defar corrective action without diverting attetvio from primary tanks. However, as with a vWualo noise and auditory load may not provide an eviraoment conducive to the detection and weogaltion of aural caution or -warning signali The funal mi of
visual and auditory caution/warning sig•als AW be based on a human factors analysis of (I) citicity, i.e., time available to respond to eaich cauti
war-
ning signal; and (2)the %isualand audio workload of each mission sagment for which a caution/warning signal is critical. The lists of warning, caution, and advisory siasals that follows are provided is a suggested baseline. Thc final eonfisfurulo su'.4 be determined from the subsystem failure modes and effects analysis, and the previously mentioned MFE analysis. tM-AI.! Warming Sl h Warning signals should include, but not be limited to, the following information: I , Engine out (identify engine if multiengine) 2. Engine fire (identify engine if multiengine) 3. Landing pear up (if retractable gar installed) 4. appliable) 5. APU Other fire fire (if Lones (as appropriate) 6. Low/k~gh rotor/ftline RPM. 6.Lwhhror/gieP.
MIL-STD-41 1.A warning signal is a signal wsemirbly indicating the existence of a hazardous condition r-quiring immediate corrective action. A caution Signal is a signal ambly indicating the existance or
10-2.4.2 Cmiee Caution signals should include, but not be limited
an impending dangerous condition requirina attention but not necemarily immediate action. An advitory signal is b signal astambly indicating safe or normal configuration, condition of performancs, operation of essential equipment, or to attract attention Qnd imlpt information for routine action purposes. Special consideration sad/ bc given to minimiig erroneous 6igaals, and to combining h, rMu input parameters through logic networks in order to provide a more credible signal for such complex and citical situations as an enine-out con.
to, the following information: I. Low transmission oil 2. Low engine oil pressure(identify engineifmultiengine) 3. Low hydraulic fluid presur (identify system) 4. High engine oil temperature (identify engine if multingine) 5, Low evgine fuel pressure (identify engain if multiengine) 6. Engine fuel pump inoperative (identify wngin if multimigine)
10-2
T
AMCP 70-202 7. Low fuel quanrtity (20-min warning) 8. Fuel filter bypass operatins (for each filter) 9., Oil filter bypw opeating (for each filter) 10. Chip detector (engine) II. Chip detector (accesory section) I1. Chip detector (transmission) 13. Chip detector (tail rotor gearbox) 14. Engine inlet icing (if applicable) IS. Other icing detectors (where appropriate) 16. Electrical system faihart (both AC and DC) I. Essential AC bus OFF 1. Main transmission oil pressure 19. Main transmission oil temperature 20. APU low oil pressure (if applicable) 21. APU high oil tumperature (if applicable) 22. APU rotor speed (low/high) (if applicable) 23. SAS failure 24. Oil cooter bypas operating 25. Low oil level for each independent oil subsytem 26. Drive system overtorque (if engine rating is sfignificantly higher than drive system),
20-24.3 Ad'Isry U S Advisory litSht should ini.lude, but not be limited to, the following information: I. AFC5 diengagc 2. Pilot heat ON 3. Parking brake ON 4. Anti-ice ON (if applicable) 5. External power ON & Starter ON 7. Rotor brAke (if applicable) 8. APU ON (it' applicable).
j•3 10-3.1
L
T INSTRUMENTS
GENERAL
Helicopter flight instruments are basically similar to those of fixed-wing aircraft although frequently they are optimized or provided with additional features to make them more suitable for helicopters, This paragraph dia.ussca preferred arrangements and other characterihaics for the selection of helicopter flignt instruments. The detail specifications will define when multiple installations art required.
/
-
\
10-3,2 AIR SPEED) INDICATORS In view of the ability of the helicopter to fly at very low airspeeds, including hover, airspeed indicators suitable for use in fixed-wing aircraft are not ac. ceptable for installation in helicopters. Instruments having increased accuracy in the low-airspeed range have been qualified, and shal be specified for heli-
copter installation. Airspeed systems that are capable of measuring and displaying both the magnitudz and direction of the relative wind should be considered when accurate relative wind data are required to improve weapon or navigation system accuracy, or when relative wind limitations during hover are critical. 1033 ALTIMETERS Barometric altimeters installed in Army helicopters shall be of the countei-drum.puinter configuration. A maximum altitude reading of 30,000 ft is adequate for use in helicopters. One of the altimeters in each helicopter shall provide encoded altitude information to the transponder compatible with tkl, CONUS air traffic control system. Sonm mission applications may require an altitude display that providts an accurate and direct indication of relative altitude or height above the terrain. In these cams, radio or radar altimeter will be specified in addition to barometric altimezeis. 10-3.4 TURN-AND-BANK INDICATORS The turn-and-bank indicator provides rate of turn, and/or needle and bail sideslip information. in some cases, the turn-and-bank indicator should be cornbined with the attitude indicator. However, the turn rate gyro and the attitude gyro shall be provid3d with independent power sources. 10-3.5 ATTITUDE INDICATOR The attitude indicator provides the pilot with a substitute for the real horizon as a rclerence for maintaining desired aircraft pitch and roll attitude under all flight conditions. A typical attitude indicator is the IND-A5-UHI. This device is an electrically driven (400 Hz) gyro that is housed in a standard 5-in. instrument case, a size that is preferred fcr ease and accuracy of reading and interpretation. Because of the magnitude of instrumentation required in m-,., Army helicopters, the basic attitude indicator #_enerally is replaced with a mo.e highly integrated display such as a vertical situation indicator (VSI). The VSI, if spocified, shall provide the following as a minimum: attitAde (pitch and roll), rate of turn, inclinometer information (slip and skid), FNI homing and station passage, glide slope, pitch and roll trim, and weak signal flag alarm. When a flight director system is provided, command bars are added to the above resuiting in an instrument referred to as an attitude director indicator (ADI). An example of an ADI which has been optimized for helicopter application, including a coll;ctive pit.h command, is shown in Fig. 10-1. 10-3
\
2J
3
II'
21
"ID
17
ROLL AT
CL.1,RT
FTR
5o PTCH TT. CALE15,
4,FI/ IECO M°FA 15 I.
GYRO FLAG
2. 3. 4. 5. 6.
ROLL ATT. SCAL~ ROLL ATT. INDEX POINTER FLIGHT DIRECTOR CMD. FLAG PITCH ATT. SCALE DECISION HEIGHT LAMP
14
7. VERTICAL DEVIATION SCALE S. VERTICAL DEVIATION POINTER
9. VEWACAL DEVIATION FLAG 10.
10-4
M-.
Typical Hldopteo
~P[YMBO
NLIOEE
13
12 11
11.
LATERAL DEVIATION SCALE
12. 13. 14. 15. 16..
RATE Of TURN SCALE RATE OF TURN POINTER INCLINOMETER -PAU SYMBOL ROLL TRIM KNOB
17. ROLL CMD. POINTER 18. COLLECTIVE CMD. POINTER
19. HELICOPTER SYM60L 20.
PITCI TRIM KNOB lm
l
CL
PITCH C4D. POINTER
AWtlti Director Iudk-atoF
LI -
AMCP ?06-202 10-3.6 RATE-OF-CLIMB INDICATORS
Typically. these mi;;mum requirements are met with
The mechanization of atet-of-climb indicStors i dTeribed in par. 9-3.4.2. Rate-of-climb indicators in-
three instruments: clock, standby magnrtic compau. and radio magnetic indicator (RMI) which displays
descibe 9-.4.. inpair Rae-o-dib inicaorsin-
sta;ed in Army helicopters shall be the rapid re"spoan accelerometer-aided type with a scale range of 6000 fpm.
104 NAVIGATIONAL TATION
INSTRUMEN-
10.4.1 GENERAL The types of navigational systems used in Army hslicoptcrs are dependent upon the mission w signed. The tynes of equipment to be installed will be defined by the detail specification for each model of helicopter. The detail specification als will indicate Government-furnished mad contractor-furnished equipment. This paragraph discusses the types of navigation instrumentation most commonly employod in Army helicopters. The navigational systems
the last throc of the given functions on an instrumentm
similar to that shown in Fig. 10-2. If mission requirements include extensive CONUS IFR flight, a simplfrid course indicator similar to that shown in Fig. 10-3 may be added to provide VOR/ILS course deviation; or the functions of the RMI, and course indicator. combined on a horizontal situation indicator (HSI) similar to that shown in Fig. 10-4. The HSI provides the additional capability of displaying the distance to the selected navigation aid or way-point. The capability to display both range 'in kilometers) and bearing to selected -aypoints stJll be provided if mission requirements include nap-of-th••narth (NOE) navigation.
2
that these dhaplays are based upon and the functions
3
performed by than are described in par. 8-3. J
.
10-41 TWYES OF INSTRUMEN'rS The instrume.,tation required for IFR flight is
X &
dzifirned in AR 95-1. The mininuimn rc-quired navigation instrumentation is: 1. Magnetic compass with current calibration card 2. Clock with sweep second hand 3. Gyrt-stsbilized heading referenceFR 4. Automatic direction finder (ADF) 5. VOR receiver (if VOR facilites are to be used).
OP
A(
01. o
Ar
02A R
Figure 10-2. RadIo Magetlk Indicator
HORIZONTAL POINTER RECIPROCAL POINTER 3. VERTICAL POINTER 4. COURSE POINTER 5. COURSE SELECTOR KNOB
Figure 1e-3. Course Imilkator 10-5
DIST
12~
CPS
0
#
vi
1. 2. 3. 4. 5.
DISTANCE READOUT COURSE ARROW HEADING BUG LUBBER LINE HEADING FLAG
7. DIGITAL COURSE b. BEARING POINTER NO. I 9. NAV FLAG 10. LONGITUDINAL DEVIATION BAR )I. HELICOPTER SYMBOL
6. LATERAL DEVIATION BAR
Fomw I. 10-6
12.
Ha•uu
BEARING PONTER NO. 2
SI..in.m Iaiw (HSI)
AMCP 706-202 &r. This approach becomes particularly attractive in
143 Map Disptys
Bemause controlling the helicopter and mainarelibopthver dmny onring th tain geographical taining geographical orientation are both very dei ianding tasks in NOE flight, a map display may be to reduce crew workload. The capability considered to display present position continuously is patticularly.y2.valuble when terrain obstacles require frequent heading changes. Map displays are generally either of the projected film or paper rol!cr type. The primary advantages of each of these types are: I. Projected Map Display Advintages; a. Simpler map preparation for a specific mission b. Larger map storage capability
c. Display can be orientcd either track-up or north up scale change Simpoe in flight d. Adantges. Roller figtsplaye Advantages, Roller Mp Map Display 2. 2. cost a. Lower initial b. Easier preflight and inflight annotation
10-5
HELICOPTER SUBSYSTEM STRUMENTATIOA
10-5.1
GENERAL
IN-
the event sufficient onboard computer capacity exists. Unless otherwise specified in the detail specification. the following subsystem parameters s/all be displayed: I. Gas generator rotor speed, calibrated in percent, for each engine, and for the APU (if applicable) Turbine gas temperature, calibrated in °C, for each engine, and for the APU (if applicable) 3. Output shaft speed, calibrated in percent, for each engine 4. Output shaft torque, calibrated in percente for each engine egni5. Total utegn torque, calibrated in percent for all engines. if multiengine 6. Rotor speed, calibrated in percent, for the main ro)tor 7. Oil temperature, calibrated in °C, for each engine 8. Oil pressure calibrated in pounds per square inch (gage), for each engine
9. Fuel quantity, calibrated in pounds, for each fuel tank 10. Total fuel quantity, calibrated in pounds, for
Subsystemn instrumentation provides cockpit rfereles that desribe the condition of enins,.eh secondary power systems, and ancillary equipment.
all fuel tanks II. Oil pressure, calibrated in pounds per square re rit_ g.rb r calibrated in 9C, for each demiferasure, 12. Oil
When pilot and copilot are Geated side-by-siue, the
drive subsystem gearbox.
most common instrunment panel arrangement groups the subsystem instruments in the center of the pantl. When the pilots are seated in tandem, it is necessary to duplicate some of the subsystem instruments and to place ther. on each pilot's panel. Preferred locations are given in MIL-STD-250. 10-$51
INSTRUMENTATION REQUIRED
Electrical, hydraulic, and pneumatic subsystem instrumentation should be based on subsystem capacity, redundancy, and failure modes. When redundant hydraulic systems or redundant generators - either of which is capable of carrying the !ntire electrical load - are provided, caution lights indicating generator failure or loss of hydraulic pressuie may prove sufficient.
Thei number and complexity of the instrLmcnts arc limited to the minimum required for safe and el-
ficient operation of the individual subsystems. The amount of instrumentation requirtd dcpends upon the size of the helicopter and the complexity of its subsystems. A single-engine, light helicopter ot.viously requires less instrumentation than a multiengine, transport helicopter. Cost, weight, and panel space savings frequently can be reslized by having several similar parameters share the same display. Since many subsystem parameters are of concern only in event of a malfunction, the appropriate parameter can be selected manually cr automatically for display when a caution light illuminates. Rapid advances in electronics may result in the cost-effective replacement of many individual instruments with a CRT display, symbol generator, and digital proces-
10-6
WEAPON SYSTEM
INSTRUMENTATION 10-6.1 GENERAL This paragraph describes the required design standards for controls and instrun.cnts for the helicopter armament subsystem. Contrary to flight and naagation displays, which frequently can be selected offthe-shelf with little or no modification, weapon system instrumentation is gencrall) uniqt.! o the nweapons mix on a specific airframe. Caa-cf, tion to mission requirements and esti fii.:" ciple5 of human engineering are required to ,. an optimized man/machine weapon system. .,e armiament controls and instruments should provide the operator wi'h rapid armamrnt subsystem status
10-7
AW;
7W0~2O__
_
j
_
_
_
__
_
104.2 DESIGN REQUIREM1ENTS For all weapon control systems, the following design rqietnshalbe ape ciosa appedtct Meulirlem ndpentsabll tote st opqertaation traoly pref equ.r frmltipe, intald
2. Two actimaions art required, preferably by two separate controls whicit should be separated so that they cannot bc actuated by one movemtent. For example, most turrets require that both an "Action" ox "Meadman" switch and a trigger be dq~ressd before the weapon will fire. Lever-lock and/or hooded conrigurations shall be used for all master arming switches. These controls shall be designed so that iradvertent activistion is prevented. A will-to-ust control catoability should he provided. Both the design and the crew responsibility for this contcol should be developed based upon human factor studies of the particular system. 3. Reversible, inflight capability of armning and returning to an unarmed, or safe tandition shall be provided for weapons and/or suspension and release mechanisms. The ARMED state shall be designed to return automaticaliy to SAFE. and the SAFE state should remain unchanged in the event. of an aircraft
mal firing or release of the ordnance. All weapon controls and circuits shall tx fail-safe so as to preve:it firing or release of ordnance in the event of improper control operations or sequences. Who-TC prauticable, the design should make it either mechanically or eleca tric llv mn~ngsMihleinP tun~,tiilt e'intorn] cr~r-z;e a - -r----improper sequence. The operator shall be ptov~dtd with feedback to indicate improper sequencing. If deemed necessary, it is permissible to have controls serving dual fuanctions, however, in no cast should safey b comromsed.Conrol usc~ fr prarn or
powert faidlur hle. eige o mu n n 4.acio Thetwieen phaowerdeined toprcludemamny cintrcinbtenpwradciia raetcr cuits. Cont~ol power shall not be applied to the wao nesi stre n tninlyb h operator of the system. 5. Jettisoning of ordnancc may be effczted individually or in multiples, provided that the warheads are in an unarmed state. Depending upon the arniantstedsgmiilaaybjtioedihr by free fall or by being tired from their launchers.
2. Where practicable, all armament controls and indicators with the exception of the fitring switch shall be grouped together. The group should be outlined hy a 3/16-in red border (Color 3116, -ED-STD-59S). Orange-yellow (Color 23538. FED-STD-395) and black (Color 27038, FED-STD-595) striped borders shall be used to outline armament groups when red compartment lighting is used. Placaro abbreviations shall be madc for each indicator and: control in accordance with MIL-STD-783 and A14A BUL 261. Preered'octirsateshwnir. MIL-STD-2ý9. 3. A complet, failure mode and effect analysis (FMEA) during the design stage is desirable ini order to preclude the inadvertent design of unsafe failure modes into the system. 10-6.2.1 4rm14, FaziLng, and Smuspension and .leue Costrol Design Require-rents for the helicopttr arming, fuzing, and suspeznsicn and release control systems include: 1. The helicopter commander shall he provided with the capability to permit and/or to prohibit prearming a~nd arming of the weapons.
18A,2.2 In -ill dividual operated order of
indication, and with rapid control of the particular systems and selection of various avak'iable options. Shape coded controls should be considered to allow the opmrtor to select the various sight modes, and types anti quentitier of ordnance with a minimum diversion from his search or tracving taks. The design sAill provide the operEtor with up-to-Ihe-minutet store inventory information in order to insure the intelligent choice of ordnance for firing. Design consideration also shall Ze given to preclude inadvertent activation of-the weapin systems.
T
_
10-8
Hism FachmsvC.nkmiewale :lose-pro'iimity control groups, the incontrols "~ be arranged so that those in jeqiamee mr in line and in their normal operation prerasakuig from left-to-right or two or nmr switefles 110om tur-iu-butieun. mu. c activated sim wiously or in a rapid wo quenc- they should be laihd so that they can be reacheo simultantstwly from a r~ position. Each control should be plarod so that it does not %inde-the operation of another control in the NqL.ý;=, With adequate clearance for a 95th percentie gloved hand. 1"4.23 Imidhess Deelp The following are weapon system indicator design lujectives: L. indicators shall have high reliability. 2. A minimum number of indicators shmu be provided in the crew compartment to show the condition (armed or safe) of critical! weapon comnponents. 3. Indic*b~or systems shall be current-limaited so as
) to prt-.ude indicator current from activating ar-/ weapon or suspension and Mease component. If wa.hwd coatinuity monitoring is required, the monitoinSg cumrnts s"Ill be limited to a value below that which will activate tie most sensitive component. Indicator circuits that are integral to control circuits canau. meet tWis reqietnemnt, and thus, should be avoided. 4. Indicator systems suadil be designed so that indi.. -r power is not available to any part of the weapon sytem unless it is turned on intentionally by the operator. 5. Indicator tests shall be possibie in flight, independent of tiii indicator-related components. 6. The operatir shall be provided with visual indication of a "hot" t-lger condition. This indication should be in tLdc operator's direct line of sight. The most conmmon indication is the use of an amber light, which alerts the operator to use caution when the weapons rmr armed. 7. The operator shall be provided with a visual or aural missile condition indication (miv~ile launched or being launched) signal. or misr 8. Immediater visual . indication of hangfires
forming weapon selection sequencing at the cornmand of the controller. The controller/programmcr unit(s) must be mounted in a readily accessible area, but not i-.oassarily in tie control pane I area. However in the case of a guidance control system for nissiles, location in the panel area may be required. The controller/programmer function may be divided between any reasonable number of subunits. The subunits may be mounted in functionally convenient locations, e.g., one per store location; or the entire functien may be handled at one or two units mounted in a central area. In alh cases, the programmer portion(s) must be easily accessible during ordnance loading. The programmer porion(s) of the unit(s) shaul provide some means of programming into the unit the type of ordnance loaded into the various stores locations. This input may be provided by switches, keypunch, patchboard, or any other suitable means. Based upon the information pro. vided by the programiffer unit(s), the controller portion(s) shall provide the function or ordnance selection, firing, and guidance (when required), and the delivery as selected by the operator. rate oforder ordnance in utimnlifif •,in renuirements, nortions
9. Identical visual indiestions arc to be employed whether live or training missiles are aboard the helicopter. This will insure that the crew acts at all times as though live missiles wae aboard. 10. Arming and fuzing indicators shall be fully automatic. Excpt for the-power-on furmtion, and the pres-to-test feature of the monitor test~ng fJncdiona, no manual operatiot, should be required. For multipie carriage, each weapon should be monitored indvidually. This may be accomplished either selectively or continuously, 1I. Weapon malfunction, if iz occurs during the presrm cycle, shall be indicated.
oi the controller which nufct not be accessible dur;..g normal usage of the system may be located wherever conenient on the aircraft. The controller shall provide an effective ground of all electrically fired ordnanct. This ground shall be lifted only during firing. Parts of the controller/programmer that must be accessible on a regular basis - e.g., rocket on stub wing of the helicopter -- may be mounted inside the leading edge of the stub wing. This would require box sizes limited to approximately 3 X 4 X 10 in. at each location.
10".3 WEAPON SELECTION CONTROLLER/PROGRAMMER
The type of insti ument to be used depends upon the condition that the instrument is recording and the required ease of interpretation on the part of the crewmernbet monitoring it. The designer shall consider: the beat type of display for the information to be provided (qualitative, quantitativc rate, trend, etc.); proper scale design to cover the required range, yet provide adequate discrimination in critical ranges; proper alphanu-meric design to assure readability under low-level illumination and in a typical helicopter vibration environment and the reliability/ maintainability features of potential designs. In addition to MIL-STD-1472 and AFSC #DH 1-3, Refs. I, 2. and i should be reviewed for additional information.
'In
5
-
zrc mumI DL pruviuc-z.
A weapon selection controller and programmer jhall be required tor helicopters canrying a .ariety of ordnance. This unit also is required for configurations incorporating selective or automatic illterval sequence launching of wistiiif and rockets from alternate xides of the helicopter. ""econtroller provides the operator with a choice of type and quantity of ordnance. If necessary, the controller als can provide controls for .iissile guidance operations. The programmcr port ,f this unit stores basic information about store a' ailability and ammunition depletion status, in addition to per-
10-7 TYPES OF !NSTRUI•1ENTS
10-9
CP7O05)201 The two types of instrument actuation are direct and remote indicating. 10.71 INTALLTIONof cversthegenralrequremntsfor MIL--597 the installation of aircraft instruments and Inoweer te vbraion ouning struentpanes, oweer te vbraion ouning struentpanes. spezifications are inappropriate for most helicopter installations and generally aye waived for the rigid mounting described in par. 10-7.2. 16-712 VIBRATION The installation of instruments requires that special attention be given to the vibration of the comnplete instrument panel. Because each instr.,ment has its own set of vibration requirements as set forth in its specification, the designer must review the applicable procurement specification for each instrument. He must determine the maximum frequency and double amplitude permissible, and must stelec a set of instruments that is compatible in this regard. Ideally, the method of establishing the vibration and test criteria for a new helicopter should be based upont Ithorough vibration analysis that defines the anticipated vibration conditions. Reference to WALSTD410 vwill aissist in this analysis, in general, and specifically for new designs, the normal procedure is to make the instrument panel as rigid as possible in order to avoid any resonatices that may be excited by rotor fundamental frequencies. (This conflicts with the vibrationi requirements of MIL-l-3997, which usually is waived.) Should the vibration charsacri~stics of an individual instrument be incompatible with the environment provided by the rigid panel, vibration isolatorr may be used in the mounting of the ctitical instrumentr.
10-10
10-7.3 ACCESSI1BILITY AND MAINTENANCE All panel-mounted initruments &WaI be mounted with the casw lugs or mounting ring against the front the panel so that the instrument wnay be installed and removed from the front. Each instrument shall be installed with enough electrical wiring or coraducting tubing to permit the instrument to be pulled out of the panel to expose its connections. Whý-re necessary, suitable means should be provided to prevet fouling or objectional interference of slack wiring or tubing when the instrument is installed in the panel. Each connection should be well identified so as to pmo dlude its being hooked up to another instrument inadvertently.
REFERENCES I1.E. J. McCormick, Humaan Factors Enghwf.ed. McGraw-Hill, NY. 1970. 2. Morgan, Cook, Chaphis, and Lund, Humanw Engineering Guide to Equipment Design. McGraw-Hill, NY, 1962. 3. Semple, Hespy, Conway, and Bus ii.'W, AawdysLf of Human Factors Data for !EJectronicFlight Display Systems. Technical Report AFFDLTR740-174. Air Force Flight Dynamics Laborntory, Wright-Patterson Air Force Bane, OH 45433, April 1971. 4. G. W. Godfrey, Principlesof Display Illumination Techniques for Aerospace Vehice Crew Stations. Aerospace Lighting Institute, P. 0. Box 19122, Columbus OH 439'19. 3. Handbook of Ae fIluminatin-P EngineeringSociety. NY. 1966.
CHAFE?. 11
AIRFRAME STRUCTURAL DESIGN 11.0l t
UISTOFSYMBlOLS =, limit Colo aceleration. number of gte gfs, dau &
i.e.. by welding. alrkarne is to be fabricated forging. buildup (riveted or bonded), or
from composte materAls - is the principal decision
g
mesioless - a mcleration due to gravity. 3 ekalero
imn-
-
a,
aeation ratio, / aoties - longitudinal aeleration, number of x"s,
diensionless2 IL VL
Ve w X
- Limit flight load factor, dimensionless - thim for accleration to reamh limit cargo acceleration. c - time to mech peak al tion, am - velocity at tune tI, fN/e - iditial Velocity, ft/e - width of the cargo floor, ft - controlled deltion, in.
| .g
1-1
A
INTRODUCTION
to be made in detail design. The trade-off criteria to be used during the investigation of fabrication techniques include weight, surface finish, stiffness and ruggedness, fatigue scWtivity, cost, and properof materials.
DESIGN CONSIDERA
ONS
11-2.1 WEIGHT .-- the begnning of the detail design effort, the weight group piovides the weight budget to th: desin group. The weight budget is based on statistical analyses and estimates of the prelimit.ary design. A state-of-the-art design, therefore, normally would meet the weight allowance, while advanced design techniques and new material- should produce a structure weigahing ic than the aiiotment. incou-poration of new design techniques in order to save
weight must be considered in conjunction with the
Structural considerations for helicopter major "oupo•net design were discussed in Captcr i. This
other requirements sioo-, for instance, advanced designs may increaw the airframe coat. In any case, a
"chapteris concerned with the detail structural design
strict accounting of weight with respect to the budget mug be maintained throughout detail design.
or the airframe only; i.e., fuselag tail boom, stabilisers, fins, and auxliary lifting wings. Secondary stucture, such as door, cowlings, and fairings, is inciuded. Detail design of transparent &ram also is di*caumsod. !bebhask. bop_.'.t- configuration is chosen and the external loads aii developed during the prolininary design of the heliopter. Fundamental airframe decisions, such as whether to use monocoque or usmimoucoque construction, also arc made durin preliminary design. The task in detail design is to confirm that the airframe structure designed in preliminary desgn meets the mission performance and curvivability requirerments, and that it can be de,eloped and produced within the budget cost estabtisbed for it. The bases for detail design are the helicopter detail spedfications, MIL-S-8698 and MIL-A-886 through -8871. and the design criteria defined by or developed from than documents. The detail design involves selection between alternative types of local structure by application of certain trade-off criteria. Material properties, method of fabrication, weight, and strength limitations are all eimpotant considerations in the design confirmation. edtemnation of how a particular portion of the
11-22 SURFACE SMOOTHNESS Surface smoothnes is anr.ther structural quality that affects aircraft performanc. CriteriN that limit the use of protruding fasteners, stipulate iraged treas, and dictate external contours are providod Oy the aerodynamic group. The extent to which 'nese criteria are applied depends upor the impact upon manufacturing "ost.It may be possible to trade off the additional cots of meeting these criteria through simplification (therefore, lower costs) of the structure itself. A curved surface often will be more costly, but it will be stiffer, more rugged, lighter, and less fatiguesensitive than the simpler alternative. Limiting flush fasteners to the forward 25% of the airframe aerodynamic ltngth may be an acceptable compromise that will yield the required drag reduction. !!.2. STIFFNESS AND RUGGEDNESS Airframe comprnents must have adequate stiffnes to meet stated vibration criteria. There must be adequate separation between the natural frequencies of prime modes end the exciting frequencies. This frequency separation will reduce the internal stresses
-I
W
-
Ai"
caused by amplification (a phenomenon cxperimced
design stage. Mitarials, tooling, labor, quality con-
Yheu operating near resonance). Vibration levels at the pilot and crew positions oftzn are excessive (uncomfortable and distracting) 4",- to poor airframe design; rotor vibmr.ion has been amplified by structurl elements whose natural frequency is too close to the operating frequencics or to multiples thereof. Vibration of structural components can cause audible noise or pilot fatigue. In addition to crew oiscomfort, vibration can cause structural fatigue and possible catastrophic failure. Ruggedness is a quality that prevents denting or puncturing of structure by rotor-induced debris, ground handling, erosion, or brush. It is difficult to prescribe physical characteristics that would prevent these types of damage. Flight testing of a prototype should include simulated operational conditions so that ruggedness can be checked. These tests will show those areas that must be strengthened, resulting in the minimum weight increment to obtain the required capability.
trol, and facility costs arc irvolved. Mak--or-buy do. cisions may influence detail design of components. For example, a fairing might be manufactured in hou&c if molded from Fibglasu; but if pressed from aluminum, it might have to be developed outside. In such an instance, weight, stiffness, etc., atso must be considered when the method ef manufacture is being selected. Table Il-I indicates the cost impact of various detail design alternatives. Each manufacturer can p1pare mch a table dofing preliminary design and update it during detail design using specific cost data. The tabl s a particilarly useful design tool if it is itated on a cost-per-air. frame-pound basis. Operating costs for airframe srictures consist of repair, maintenance, and replaomat parts costs. The design objectve should be a maintenance-free life equal to the anticipated service life of the helicopter.
11-2A FATIGUE SENSITIViTY tatýi-ally utiu. iigbweighi deigns may be
11-2.6 MATERIALS In preliminary design, material selection has been compi-ited. A most important detail design conca-
tigue-sensitive. Attention must be paid to the detai's
sideration is material verification. A wide variety of
of the airframe structure to prevent fatigue sensitivity. For example, designers would like to eliminate the clips attaching the stringers to frames in the fuselage, thus saving both weight and cost. However, experience has shown that skin and bulkhead flange cracks occur when the clips ame omitted. A highly loaded airframe fitting may have sufficient static strength because a high heat treatment is provided, but, as a result, the maturial may be notchsensitive and prone to fatigue failure. Orain orientatio, in highly loaded fittings is important. The most efficient structural design is obtained by orienting the longitudinal grain in the di rection of the primary load. A fitting may be fatiguesensitive if the transverse grain is oriented in the prinwtry load direction. 1 . Ctribution 11.5 COST In the design of the airframe, three cost areas
ferrous metals, nonferrous metals, and nonmetallic materials is available. Characteristics of thise metals and materials are described in Chapter 2. Additionally, Table 11-2 summarizes materials, characteristics, and their uses in helicopter airframe construction.
should be considered: cost to develop, cost to manufacture, and cost to operate. The cost of development includes that of design,
radar-transparent materials and/or nonreflective coatings. Reflectivity characteristics of the key aspects of the helicopter also may be reduced, at least
11-2.7 SURVIVAI1LITY The survivability characteristics of an Army helicopter desi2n include: 'I. Detectability 2. Vulnerability to enemy ballistic threats 3. Crauhwor. hiness. Detection methods to be considered are radar, infrared (IR) radiation, acoustics, and visual. In the design of the airframe structure, the greatest conto roduciig the detectability can be made by reducing the radar cross-section of the holicopter fuselage. The techniques available include the use of
development :f new methods, and testing rcquired to
for selected radar frequencies, by careful attention to
prove the design. Dsvelopment testing is a very important design tool. A number of design alternatives can be tested under identical conditions to find the best design. The coats involved may be high, particularly where fatigue testing is involved.
the shape of the target presented to the transmitted beam. The specific ballistic threats to which the helicopter will be exposed and the desired levels of protection will be stated by the helicopter system specification.
Manufacturuig costs must be estimatc'! during the
The use of armor materials to defeat ballistic threats
11-2
)
~AMCP 706-202
TABLE si-I. COST IMPACT, AIRFRAME DETAIL DESIGN COST AREA
WELDED
RIVETED
CAST
FORGED
BONDED
LOW NONE NONE
MEDIUM NONE LOW
HIGH LOW HIGH
HIGH MEDIUM MEDIUM
MEDIUM MEDIUM MEDIUM
MANUFACTURING MATERIAL TOOLING LABOR UALITY CONTROL ACILITIES
LOW MEDIUM HIGH HIGH LOW
LOW MEOIUM MEDIUM LOW LOW
LOW MEDIUM "* HIGH MEDIUM
LOW HIGH MEDIUM HIGH
MEDIUM* HIGH MEDIUM HIGH HIGH
PROOF OF ADEQUACY ANALYSIS GROUND TEST
LOW NONE
MEDIUM HIGH
LOW HIGH
LOW LOW
MEDIUM MEDIUM
DEVELOPMENT DESIGN METHODS TESTING
*IF BORON FILAMENTS OR THE LIKE ARE USED, THIS COULD BE HIGH.
"*DEPENDS ON QUANTITY--HIGH PRODUCTION RATE RESULTS IN LOWER COST.
.9 TABLE 11-2. MATERIAL SELECTION - AIRFRAME DESIGN
rWEIGHT,
lb
STIFFNESS psi
3
I1
TENSILE STRENGTH psi
COST S In
AIRFRAME USE
FERROUS METALS: CARBsON STEELS
ALLOY SIEELS
STAINLESS STEEL
IAPPROXIMATELY
30 x 10G
0.3
-
PH STEELS
180.000 300,000
LOW MEDIUM MEDIUM HIGH HIGH
MARAGING STEELS
NON-FERROUS METALS. ALUMINUM
0.1
MAtNEbIUM TIIAriUM
0.06 0:15
t
0.03
I
i5,000 30.000
10 x 105 6 x 106 5 x 106_
LOW MEDIUM HIGH
FITTINGS FiREWALLS FITTINGS SHEET ELEMENTS SANDWICH FAC!NGS
CASTINGS
FORGINGS
NON-METALLIC: * THERMPL ASTICS IHEFMOSETTING ELASTOMERIC
GLASS LAMINATES FIBERGLAS
5000
1LOW
LOW
j
5Lu
IiWORK 0.6A6
GRAPHITE
I BORON I *WHERE OPTICAL QUALITY ISHIGH
LOW LOW
HIGH'
TRIM
FAIRINGS WEATHER STRIP
I GLAZING PLATFORMS DOORS 60-400 FAIRINGS I COMPLE1 E STRUCI URES 11-3
is discussed in Chapter 14. In the diesig of airfirame structre, care must be taken to minimize the ps-billty that a single hit by the stated threat - including explosve and/or tucendiary prjojctle when so specified _ will cause a cauk, crash fire, in-flight (Ire, or comparably catastrophic result. In the design of structure surrounding Pool taks it Isessential that redundant loads paths be provided. The principal load1-carrying members may be damaged by bydraullic ram effect following projectil impact, Particularly if the hit it by a high explosive or armorriercing incendiary (HEJ or API) projectile. Ref. I provides additional design guidance for the rm duction of vulnerabilift, to ballistic threats. Regrrdleas of the intensity of combat or the soverity of the ballistic tret(s) to which a belacopter will be exposed, the design must be crashworthy. Craithworthiness design criteria for Army aircraft are given by MI-L-STD-1290. Additional guidance and specIfc design techniques for meating the stated crnteria arm provided by Ref. 2, Some of the crashworthiness considerations applicable, particuilarly to the aitfranie structure, include: 1. The inconiaoratioipi of crushable structure out. side the occupied zones to assist in the absorption of impact energy while Maintaining a Protective shl 2. The incorporat:on of turnover structure &de quate to maintain the integity of the protective shell following impaict with the ground in either a rolled (90 deg) or inverted (180 deg) attitude 3. The provision of support for the main trans. mission WWi Rotor mast so that the transanissica is not displaced into the the protective shell following either specified crash conditions or the strike of a rigid object by the main rotor W._
propellers. flight controls, electrical system, avionics hydraulics and pneumatic&as, -ntuissln landlz'g gear, crew stations (flunishinge and equipmIn) armainent, armor, and protetve devices. Weight and balance - as wall as operational and flantiousl (aotore - ane to be considered while locating the sub.. systems. T7he mounting points of each subsytaim are arrngepd sn that the overall load distrbution system of tho structure can accept the local load food-in momt efficiently. Shelves, beams. frames, mud bulkheads are types of local structures used to take the mounting hard-point or fitting loads and to distribute those loads into the basic structure. Mathods of locating and aittcIn fittings, supports, frames, bulkheads, skin and koqeron systsems, corrueion protection, and electrical bonding are major design requirements to be considered while accommodating the subsystems in the airfrarie and applyiDg the trade-off criteria discussed in per. lI1-1. Fig. 1l-1 ilustrates some of theme structural comaponents. During airframe detail design, all areas subject to repeated high loads should be anayzed for the inoornoraiion of fail-safe fasturm- Trhis desio nhi"in sophy - to prevent catastrophic failure - requzips provision of redundant lead paths so that, if one pt fails due to facigue. the load is carwried by diroreaming structure until the noxt inspection uncoor the failure.
,6A W6a~sa my1
BULKHEAD II
LNEO
1thU
propriate with restraints adeoquate to retain the crew and other occupants within tht protective shell following specified crash con,;itions 5. The provision of emergency exits of sufficientFRM number and sine to permit the evacuation of all occupants in the minimal time av'ailable regardless of the postcrash position of the helicopter. Crashworthiness considerations pertinent to the cargo compartment are discussed in par. 1V4.2.
SRNE
11-3 DESIGN AND CONSTRUCTION During detai design, overall layrouts of the airframe ame prepared from the prelimiinary design data available. 2Firi locations of the ma'ior subsystems te be supported by the airframe are shown, together with directly associated assemblies such as fins, stabilizers, and ttub wings. Major subsystems ame the power piant, transinisaions and drives, rotors and 11-4
SPOT SL
Fkgur I I-1. Airframse Ceupseest
AMWPO.0 11-3.1 FTfil GS Fittings provid, the structure! transition between two dlffgt typeo of structure, and also serv as a amvenient point for disassembly. They may be dsisagrd as weldeid, cast, machined, or forged substructures. Appeopuiate safety factors must be used wham calculating fitting strength. For exanple. castings raquire a&aytical safety factors ranging from 1.33 to 2.0, depending upon the quality control &tandards to be applied. Safty factors are governed by the casting clasaificAtions in MIL-C-6021. The decimion as to type o'l fitting depends upon weight, production quantity (or cost), and the nature of the loads to be transmitted. Castings generally have low elongation c sacgetics, and, therefore, have poor fatigue dircbmctklst Metictlous camv should be given to keepin# the local eccentricities in the fittings to a minimum. All chatiMe in crows section should be made as gradual as possible by using generus fillet radii. Abrupt chaonge in crass section cause stross concentrations, antherefore, must be avoided. Lug analysis requi-e special consideration. Refs. 3 and 4 contain aiscussions of this subject. 4 Snof I -F mr~o .rnteenra.smch na bthi rivet% bolts, for a sirgle fittling attachment should be avoided. Rtivets fili the holes and. therefore, pick up load befon' bolts do. It isimportant to preload bolts so that dlamping oi the facing surfaces is acPmnplishe and bending of the components is minimizod.
will be ovideni, and, again, cracking will occur afte a period of service. Figs. 11-2(8) and (C)show correct and in%;orrect deaign concepts for bracket &UaGhments. 1 RA S Frames arm used to reduce skin panel sine and striulger column length, to maintain aircraft contou-, to transfer various loa loade to the outer skin as shear loads. Fig. 11-3 shows an example of a frame reacting shear flow from a tank liner. The weight of the tank and the fuel result in a bending load on the liner, and the resultanit shear flow in the liner is op. posite in direction to that shown acting on the frame. The; frame flange, that support the skin must be suafficiently thick to prevent tension field wrin~kles from propagatin3 when the panelis mre not shearresistant. The moments of inertia of the cross sections of the frames must be adequate to restrain the stringers against column failure. Junctions buwcen fromes, skin, and s~ringers most be clipped. INCbRRECT
)and
CORRECT V
-
V SEA CENTER
v
I'N
A(A)
113.1 SUPPORZTS Shelves, beams, and brackets that support equipmet and subsystem parti generally are constructed by assembly of shee-metal components. The support must be strong cnoi~g to takc the design load -- the weight of the item supported times the design load factor - plus the applicable mechanical reaction forces and the vibratory loads. The loads being distributed into the primary structure from the support should not induce secondary strease, which can cause the primary structure to fail. Secondary loads arm caused by tht. dflection of the support. Deamis having an open section shojuld be investigated for shear center location. When a beam is loade off its shear center, twisting will occur, as in"diaed in Fig. 11-2(A). Secondary loading and lower subsystem natural frequency generally result.
(B)
CRACKS
food-t the supoa to dcarod.esoitnedss ina Brackts shupo iorfoed-out is parallel toasheet-metal face. If the load is perpedicular, local bending will occur and subseq~uent cracking will result. Radii should not be
count-
tocryla.Sfnsi upo
h
upr
'
'N.I
f IC)rsc (~ao1-.Cwct
igr
12
o
d&
serc
A
itui
1-
*.17
OUTER SKINDEALOTRSI
CLIPS
'1FRAME
VIEW A
STRINGER.
LINER
RIVETS
STRINGER IN
RIVETS (TYPICAL)
f
010
-
FRAME FLANGES DETAIL VIEW A Figin Ila3. Frinme Used for Ttuk Suppor
11-3'
]BULKHEADS(
Bulkheads function much the same as frames, but have udded capability and utility. Generally, major loads - such as landing loads - ame introduced at bulkheads and wre redistributed within their own plane. Out of-plane restraints, such as fore or aft fuel tank [Qtds, also may be provided. Stiffenes in the plane of the bulkhead distribute ritting loads into the web. The stiffcuners must have sulliaenat inertia to stilfcn the web against catastrophic buckling. The webs of all bulkheads should be designed to prevent oil-canning. Bulkhead .¶langes must be stiii enough to resist coiumn buckiing, if so loaded, as well as to provide skin restraint. Loads from fittings should be sheared into the bulkhead with as little occ~tricity as possihle, Edge distances of fasteners should not be less than two times the fastener diameter wherever possible, Low-strength blind fasteners should be avoided. Cutout* should be reinforced, while attachments to an unsupported web should not be madec because vibration may cause cracking. The strength required for niajor bulkheads usually is attained by employing forgings. Final dimensions are obtained by machining. 11-3.5 SKIN SUDSYSTIEMS Skin subsystems arz dcfied as structural covernugs composed of longerons, stringers, and skin. If sandwich conittruction is used in a monocoque conflgu-
[MUun. someC 01 tCSC =G5weniW
weE not usea.
Longerons ar6 the primary aitial-load-carrying members, and are situated at optimal distances from the bending neutral aixes of the airfrmniv. They amr de. signed to withstand column loads and the secondary effects of skin tension field. it usually is nacessary to locate intermediate siringers between longerons to coatrol skin panel size. Bending material (stringers) should be distributed among a large number of elements. in order to limit the reduction of static strength following damage to ons. member. Again, oil-canning should be provente by proper panel sizing. Fig. 11-4 presents suggtsted panel dimensions. Stringer and longeron splices should be scarfed and multiple attachments provided. The fasteners should be loaded in shear. Fastene types should not be mixed in a splice. Doublers should be providod around exces h*oles through the skin. Because repeated ren, oval enlarges the fastener holes, accss hole covers should not be designmd to be load-cprrying. Minimizir4 the number of cover fasteners makes servicing faster and, hgonor, more economical. 11-3.6 CORROSION PROThCrION Protection of parts against corrosion can be accomuplished in two wayt. One method is to app~ly coatings to prevent corrosive atemospheres from cowning in contact with airframe parts. The other isto in-
AMCP 7*2 2P
ALUM INUM /.LLOY PANEL SIZES FOR SKIN (WITHOUT A0 PRE-/ CIABLE STRESS) EXPOSED 10 AIR
16 -UNSUPPORTED
10
R/t. 1500
1
~12--4R/t
=
87
00"00
.(FLAT)
_-7THE
Q 5DOR/ 31
USE OF THIS FIGURE IS SUGGESTED IN OBTAINING SICIN -PANEL SIZE COMBINATIONS WHICH WILL BE FREE
-J
EXCESSIVE VIBRATION. DEFORMA-
00el 4FROM
TION OR OIL-CANNING .VALUES ABOVE THE PERTINENT CURVE WILL BE CONSI DERED QUEST ION~ABLE. R=RADIUS OF CURVATURE OF SKIN SHEET. in. ttSKiNTHVICKNESS =
-
4
/
0.04 SKIN THICKNESS I, in.
.Z0.03
UU11
I
000
0.05
0.0
WHEN PANEL LENGTH EXCEEDS 3.0 x WIDTH, DIVIDE LENGTH BY 3 TO OBTAIN WIDTH FOR USE IN DETERNINING THICKNESS FRO)M CHART.
-
Figur 11-4. Ahamahme AN"a Ps"l Sims auce an oxide coating that prevets further oxidation. Many coatings can be applied, ranging from zinc
*
ch~rnMate!
J
primer (hr internal
em-rfarot
tn
nocessary to product electrical and clectrcnic system installations having a'ccptably low levels of electromagnetic interference (EM I).
enn w
paints. Par. 2-6 provides detailed information about 11-4 CARGO COMPAR'rMENT paints and finishe s, welk as about special processes. Cargo compartments pose spccial design requiresuch as anodizing for inducing oxidation on alumiments in th~at cargo floor and tiodown fittings must be num parts. dsge oacp ttcadcahlas dsge oacp ttcadcahlas Treatmvents such as anodizing on aluminum, and cadmium- and nickel-plating on tteel, have de. 141SAI OD 1.. TTCLA) leterious effect upon the fatigue strength of the Static limit fl~ght loads sAud be used as tt.;. basis for metal. Partse that have been so treated and arc subdesign of all structure and fittings within the cargo ject to alternating loadings should be fatigue-tested. 11-37 EECTICALBONINGcompartment. The primary areas of concern are the 11-37 EECTICALDONINGcargo floor and the tiedown fittings. For helicopters having cargo transport as a priIn order to provide a continuous ground throughout the airframe structure so that remote electrical mary mission, the basic cargo floor should be designod fow a limit-flight-load floor pressure oC 00 componenis may be grounded adequately, nonpsf , * here %., is the maximum flight limit load falc-. conducting structures must have aconductanc bond tor. For light utility helicopters which are not pribuilt in. manily cargo transports, the lintit-flght-load floor Par. 7-6 includes detailed bonding requirements 11-7
Wpressure ca
b moa t
,tpaf over that por-
load under constant pressure. Solid rubber tire or
metal wheels do not offer this flabWty aomwdiglion of the f'oor that can be uwd for cargo. Baggage compartment floors should be designe for a limit- ly, oioncentra ted pressure loads applied to the Ame by load pressure of 1Oibi pef these wheels must be assumed to In, -- i proCargo tisnapoit helicopees havisg the capacity for portion to Cie flight load factor. 111dmrvaisisii of floring wnder solid-wheel loadin shul be 4a. Vehicular loads must ba designe fo the o011n10111trated loads applied by the %Vbwe whees. ThU= urmilnd experimentally, with pertmms opWOllega wheel lWads shoiuld be assumed to be actin at my Iimutltioe Spec11fied. In Most otaPormy hell-- roadin.dsd o copters, the use of shoring planks is point within the treaderay area shown in K&g 11-5. The dimensioneshiwn on the drawing describ the protect the floor from solid-wheel deaualls, suggested minimum treadway area. MiL-A488S specifies weer test requiremesats for The portions of the treadway dthatre used only for newly deuignad flooring. Nontreadway arns also ground operatins (loading and uloadng should be should have the capability of carrying limited mondesigned using a limit-loaid faorm of 1.0 applied to centrated loads in exom of '30?htý pef. A load the mA xiinur wheel load. The ara of the treadway strength of 1000w,,, pet typci y~ usdin the upon 'vbich the vehicle rests in fligh. must be do. design of cargo aircraft. The extent of the Ifflow -, signed to carry the maximum wheel load multiplied per area of application is limited by weogh andbel by the maximumun flight 11mit load'factor x,. ance considerations and running-load limits, which Maximum wheel load should be equal to 1/3 of the interact. maximum anticipated vehicle weight. This assumes a The running-load limit should be established oy laterally symmetrical vehicle with 2/3 of the weight the prester of the following: twice, the maximum carried by one axle. The maximum wheel load, multi- wheel load (i.c., the axle losd). or 30kb wIbi-inei ft. plied by the apprcprite limit-load factors will be where w is the width of the cargo floor In feet. The applietd to th.e utradwefy aesthrough pneumstic irunnin:-load limit is the ina-zismum load thast czan be tires inflated to a maximum pressure of 100 psi. This applied to any single foot of floor length. locl pessreof 100 psi can be assumed te. remain Floor tiedown fittings should be available in suconstant with the application of helicopter flisht ficient quantity and capacity to restrain the mauilimit-load factors, because the action of the pneu- mum design cargo weight under the ultimate crash matic tire under moderaey Increased load is simply loads. zo enlarge the wheel contact are in proporton to the The cargo tiedown restraint factors must be donC
---RGO COMPARTMENT FLOOR
I
IiW
88 in. mnun[
AFT RAMP--.I
~4 in.max LOORAND AMPCONCNTRAED LAD lO BASIC CARGO LOADING RAMP L -J TREADWAY
U
CARGO FLOOR TREAD WAY
jUN IFORM STRENGTH
300n,,,
jPRESSURE
FLOOps
psi
LOAD 100 psi CONCENTRATED LOAD (n-1m/3),,(MAX VEHICLE WEIGHT) PRESSURE LOAD 100 psi CONCENTRATED) LOAD (n., ... /3)x (MAX VEHICLE WEIGHT)
Figuire 114. Mhmimeva Fleer Strength Requiremmilus
-
-
-u
So&ash nu w hewhis dopie lduqmla Im~ pbcaIq ib ca spu§ ine tMbeap flie fr th. SXW, W&S, N&u. and 95Speram" usuua thel smiaC. looedown011 wueb" anile mmnhqbls helopee sadskas, and an shwn deemed Is WL-A4SS. in ftg 114M3. ltlama be sum that, In .,tt to fletasaofsaewul agmqeulflcplsuofmwpto 'hims cargo rbI* bo the helhepes- fleow throughreIs hM nahw dom stlnfled hut hateors a 941-pemetils helicpte amok It would be cmt qmodhlty of the user. The dul4f-t' rmpm"b* okmeeu to employ wme thiem the mushe of1U is to critw dude's fiudimm thea will carry ibels downs reqaku hr the -wa (fth pervnemil) wash. SWtkg issaled maud bad withou Mar. Thew This beomese awoMblbinv Unis of weih and cornshoul swivel aid rotat through a hemluphen pietRy. bodaded by the Ulcer. The rated load shoul be corn3at efforts in the developmentn of loadl-Ulvltng da& u-1es the S sie lead for the purpos of docarg tedowam hae shown thee devicse to be -n oSpins the Kiting diSts supporing strauctr For 'Wsti solution to the ash-load realmsn problem that thnsandard h reammued nis kqp hehepe ksa, (Ref. 3). Load-limt~n iedom are emegy-mbeorbý fero tiedowa Suling be rated at SOW) lb ultimate lag devic that link carg to the ealeag heliopter str%*t. Tbs rmw of fittinans use stedl of the floor tiedown.. They are designe to reds - fitfloor shouldl be rated at G0,0W lb each. This facilicreasing tensio, load rigidly antiL, at a proI.t with aminimum tan the rmuetmat of vWhml ar determilned force, they yied and deflect owt some of deftwa as.dMnf while malnt~lnl§a co~tnsistan distane W1= W" MS WhAMA onnctedin ~w viththecargo tiedown chains, such devices &ac to ettenuste the acIt baa bou found that cupg compartienmt boW-ll colorations tranmItted from the heLicopter floor to twdlai awiulallom can exceed 27 £ during a mtiotc the cargo, as sboi. In Fig. 11.7(A). vivabi crash Wrdlag. FIg. I1lAFA) shows a typical
-
.
ý1ýr 11"a"UM \
rmi
SOW i. a i,1kw - iw
saum
UW
vWftU3-&%
%4&aa -%mfa
cargo and the airframe as they act separately during H Considerable; eperineatal work in thi arehabaaIRR shuwn that the typica major impact accelerations opurpoesby aslut-CAG can be rpreseen~td lot ngnmSug during thetat cavsh of en iauslrwnenled helicoptar
} (WM. 2).
i
A
TWICAL MEAIJMD
EATS PC~
SUGGESTED
TMANQJLAA
L
DEIN U~iq
AVERC
SL~L 0
~~~~~~2t,
i* TIME t.
JA) TWI CAL MA.AQR WACT ACCtIREATI'JN4 PULSE
2
-
S1PERCS4TILEUt.-0.Oe.
9ftOPER#iT1LEjc,-C.1O4).
10
0-
INseeM
V0
t
:rr
IlR
HSG
IN-W
M
rA
VL*
IAME CA11RGO
FeinEi,-om
H-13
,a0240). 1TL, H-4
t
YO t, seIMc.
1C
A
O
DCR
sm
020
H2?
()Ac.AlNT1
oedt Acemlaralle Fee ao I1r
71"
e 11-7. iffis of Lae
1 ir Wed.
ELi
the crabh event. Thee curves are derived by into. gration of dme respective so lration curves in Filg. I l.7(A)ý %cause it is wnder a lame aocalcratio, the cargo mast mow*eslightly farther than the fuselag during the period of the aocce;%eration pulse. This additionall diett-ace is the distance through wvhich the load-limiter must deflect. and is determined by intograting the two volocity-time curves (cargo and airframe) to find the distance each-travels during the event, and :ban finding the differenc between the distance.. Eq. 11-1 can he used by the designer to do. ternune load-limiter stroke required, assuming a trianguwar acceleration input pulse. 2 K ~A? f (11-1) 9~oeg:24 2rk-~,I 2 KIt where Hf - pulse peak acceletration, number of S's, dimensionicis g - acceleration due to gravity, 32.2 ft/sec2 - acceleration ratio aL/ H. dimensionless K 4L - limit cargo acceleration, number of g's, dimonsionless to - time to meach peak duration, se
)90th-ercentile
(_
CLASS A CARGO
U
U
This equation is derived from Ref. S. The WOa-limiting cargo restraint concept his beon developed to the prototype stage. Coat and weight permitting. it is recomenwded that conslderasion be given to incorporating woci devioee into the basoc dessign of new helicopters. Such loed-lsimiting mo straint fittings would provide a maximtum degree of crashworthinms. Details of load-limiter desig and application ame available in Ref. 6. When speciIfring loadl-limiters, the "Ail~erg must make a trade-off between rated streengto and loadlimiter stroke. Fig. 114 shows the various ways of achieving adequate restraint of Class A cargo in a crash, using load-limiters having different energy absorption chiaracteristice. can be soon by rsforrinS to the sample curves that an additional resariczion on the performance of the load-limiter is that the load-delctin curve must avoid the shaded area below the base curve. Thiq bass curve indicates the load below which the load-imiter must act as a rigid link in tho tiedown chain. This load should define the rated yield strongth of the limiter. An optimum load-limiter, from the standpoint of weight and number required, would be ane
X
ffI-II ~IL
0
1-0
LOE 10
SE 15UR)V5E3 CONTROLLE DEFLECTIONCURVin.
CNFOLDDFETO
)
AMCP 706-202
with a load-deflection curve conforming as closely as posaible to the bane curve. The long stroke of such a limiter, however, could allow the cargo to move sin ecessigive distanca within the compartment. It is recommended in Ref. 5 that Class B cargo (vehicles, etc.) be restrained to withstitnu the 80thpercentile crash pulse of 13 g, rather than th.e 90thp~rcenitile pulse that defines the restraint criteria for Class A cargo. The rationale is that personnel rarely are carried in the cargo compartment with Class B cargo, and, thscre'orc, that only the relatively remoie flight cme need be protected from shifting Class B cafgo. Fig. 11-9 shows the load-deflection requiraments pertaining to load-limiters restraining Clas B cargo.
11-4 I
~the
TRANSPARENT AREAS
The characteristics of available glazing materials are reviewed in par. 2-3.5. Additional information on the properties of the matorials are available in Part 11, MIL-HDBK-17. The optical quality of the windshield must be maintained under all conditions of loading. To keep opticai distortion to a minimum. flat panels should be used to the maximum extent practicable. In any case, considicration must be given to the deflection of the windshield under load and to the effect of these deflections upon the optical characteristics. The loadings that result from thermal gradients acrosb the windshield must be considered :,eparately as well as in coin jination with pressurc loadings in order to dcttrimine the critical loading condition. The effect of
an antireflective coating used to reduce the glint from the windshield and canopy must be jinduded in deter-
The design requirements shdl be as established by helicopter system specification -Two categories of transparent areas are apparent. One category is represented by a windshield that must i.;,hstand direct airstream loadb, possibly including impact by birds. The other category is represented by an ob=-mo.wn-- !=.*a n.....ie, 1---f or floor of the crew or passenger compartment. Such a win-
mining the thermal loading. Ijistahation of s'lazing -materialsis described in par.
dow is subject te indirect pressure loading. positive or negative.
support from the window. This is pi.rricularly important for window openings within doors. The dc-
2-3.5. ?~art 11. MIL-HDBK-17 also contains a thorough discussion of the design of thc "edge attachments", or means of fastening a Slazing material to an airframet. With regard to the structure, it is r cmmended thatt on.-.inac fnr trnt~p ro-rt pr-sh self'-supporting, i.e., not dependent on any structural
CLASS B CARGO
fXCNRLE..f~C!
-LJ LUL C-,
CA", -IL
~LOWE R (BASE) CURVyE
)CONTROLLED
DEFLECTION X, in. Figur 11-9. Class B Cargo Forward L~ad-deleion Eniedop
AZCP ?06-202 tolerances chosen will bec the result of many conarderations. TMcac include. 1. Number of units. The total number of helicopters to be produced influences the type nod extent of the tooling that can be used. A large produrflion run will permit intricate tooling with co.-;ts that can be spread over mony units. 2. Fabrication method. The way the helicopter wrill falnicatod on the production line influences the 11-6 EVELO~Mbe tolerancoes. If thc helicopter is fabiicul-d frim many The iterative promca of desijin. buildk, test, and recomponents that have been subcontracted to manay detigia is called developmenL Several different verdifferent venodrs, there is a KIeed for closer tolerances thnitehecoersetrlyabiednoe sions of the same airframe component mray be desigaed, built, and tested to determino. trade-off pamaproducton line. r eters such ua weight, streng~th, and cost. With thene 3. Assem bly method. There are various mnethods; data avaiiable, a decision can be made as to which of assembly that have different tolerance restructure bwAL fits the requirements. Clowe tolerances will poi mit complfte inDevelopment, includir.g testing of alternative terchangeability of parts without additional work. manufacturing methods, is a costly process. It should Selectiyi: assembly will permit less strict tolerances. A be used onily for redundant strtceturej-, where unalysis large prodtwtior. run .. ill make parts svailabic for acletvmacigwhadcntpr.Ifasby is eixher miore costly or is impossible to accomplish methods and production time is available, the saw to withn schdul, te aloteo fornewconsrucbon suit, file to fit, approach mnay be the best. xpe~ucuni i 1~hch n~iufa~lur asmetechiqus quired to prove feasibility and to dewirmine costs. wfoNrmarnc. The nood for flushness and gap For purposes of this discussion, devclopnvint does '-:~ to M. C,-----------i6uimi wil be not inclu-e the testing of oarUs to detvriniijii points of an influcnc on production tolerances. k(,cal failure a.-d the changes necessary to obtain atis.. S. Interchanicabil-ty ance replaocability. The extent that parts e!nd components should be interstrength. This testing, rework, and rc testing is considered a part of the denionstratinn of structural changeable and how they Ls.%uld be replaceable have adequacy aind is discussed in par. 1-8. a direct influzznce on production tolerances. Replactritnt requirenients at the lower maintenance echelons will require tolesances that will permit reMANUF~ACTURE placement with a minimum amount ef match fitting. The d~igner has the responsibility to provide 6. Cost. Cmst production tolerances who-e they are as is cost ufactur~ing man a as low requiring airframe not required result in added expense. Loose a~s esil maufctued posibi. ie enceil, po.irfma Gencaflso i theleastst eslyone. A rasonab tolerances that present assembly problems Also may ~ ~ noocpne number of subassemblies sho&*d be planned to allow consideration must be given to production easeof nnuacue.Careful The final tolerances used will he the result tolerances. reusually structure of the areas all Accessibility to of coordination between the procuring activity, su~s os. bcauc a loer mre orerscanbe pperod f cnstuctoa.engineering, manufa-,turing, and purchasing. I a&vs pled urig eChpr ,AMP7621foprlmny The structure should be designed to permit use of design treatmncrt of accesuibility/interchangeautomated machines egriveters) and numerically abiy/elcblt. controlicd machine tools during manufacture in 14 SBTNiTO order ta minimize production costs. In Chapter 4, AMCP 706-201, the consideration lIn addition to design, development, and manufor practical prodmaction tolerances is discussed. The facture of an airframe, it is necessary to demonstrate prelimianary design will have coilsidcred the that the structural subsyste m neet the design ~okerPa-cs that can be achieved during mnanufacture requiremnent,- Two methods are used to demonstrate and will have been the basis for design selection. The the structural adequacy of the airframe: analy-fis and example used in Chapter 4, AMCP 706-201,* is the testing. Generally, it is acceptable to use one or the pihlimirnary design of the doors and hatcnes. other method to prove that strength and deDur; ng de!.ail design the exact prodtction formation, utility, dynamics. and weight and CG are tolerances required mo~st be established. The flections of the opening should not be large enough to case5 failust~ of the glazei-upporting systean. Thickntsa of tht gisawx determined by analysis or test - should be adequate to support the pressure and impact toadings, In addition. the glaze mnaterial must satisfy any ballistio-resistance requirements containod in the Fvstem itpecificatior.
.quireinents.
4.qwae
I11-7 #W.
11-12
9 ducs an optimum design that will be proven *dequate by test. When both analysis and testing are used, significarnt savings in weajht can be obtained with designs 11-8.1 ANALYSIS for which the analysis indicates a small negative or =eo margin. During testing, thte areas in which data As the deveclopment progresses, substantiating arm prepared in acros danoe with the contract data rt- failures occur cani be strengthened. Sonme of these areas will have sufficitot strength as it. In this way, quirernents list (CDRL). Specific data requirements are ooordinated with the Airworthiness Quali- the minimum weight can ýe achieved. Sonmc portions of the airframe sme designed with fication Specification (AQS) prepared and approved load paths. In the event that one path fails, multiple possible for an individual helicopter program. The scope of these requirements is discussed in Chapter 4 thc others will continue to carry expected kvads. It of AMCP 706.203. Chapter 9, AM.CP 706-20-4. pro- woil be necessary to tust these areas with selected load senits a comprehensive discussior' of the final quali- paths failed, and with thti applicable fatigue spectrum applied, to demonmtrate that the structure will confication of the airframe. The analysis and testing performed during the design phase are discrv zed in this tinue to support all loads uxrti' the next mandatory inspection. It also is necessary that inspection of these paragrapl. the start of the detail design effort, desifn pram- areas be possible, and particularly that portions of necters established during the pretimiraary design crw Atriiturc thatt are designed as fail-safe not be obixarmCd. Chapters 7, 8, and 9, AMC? 706-203, desconfirmut or amended. Tht loads applicable to #,Iairframe components undcr design Wlglat and grow~d cribc component tsts, surveys, and demonstrations loading conditions wmur be established iind the Oth Piro. arptlicabic. Fvtiguc testiing may be generalived in that only critical desian oondiu-_ns deterrmirad. In Chaipter 4, 4etfl~i~ (Bitch as zi-1) arc collect~ed. Statistical &;-; 7OC-201, the dcsixra loalding CoLdizns described and procedures for IL.ý dett-inimti.n oi~anilyses that consider operating load frcq~ienc) will temnal load distributions are discacu.sd. Leo'a com.- minsre. ae~quatc structural reliability. The number of iimducted will determine the confidence level r.4 Cit prehensive compul .r prograim su.ch at NASIRA.N foi- tht dcsign. Fatigue life determination is dis(Ref. 7) is reconini ceed. cussed iti detail in Chapter 4, AMCP 706-201, and the Following determination of thek- a(eings thal arc riiquirtJ tests are outlined in Chapters 7 and 8, critical for individual parts, in nmny ciucs strcvtgarcý AMCP 706-201. adequacy can be tubeta;'tiated by stress wnelysis After the hoicopter prototype is manufactured, alone. Exaimples are parts thstt involve either a!ýirtglc flight aist loads will be gathered as basic data to be load path- or simple rc4unclarcy, for wliých thereforrc u&Wi for structural qualification analysis and testing. the stress and analysis it boib simple and accurace O~ti- er ar for whbich ansil asaalnn r s adeouste are It may bt necessary to redesign locally before qualisimplet fittings for which dynamic"('fatigue) loadings ficationi if~ the flight, loads are in excess oi those preare not significant, and secondary structure and comn- Clicted by analysis. ponents that are classified as nonstructural. However, because the airframe structure gonecrally in'REFERENCES cludes multiple load paths for which even the most sophisticated stress analysis maethds often provide I. W. D. !)otbeth, Survivability Design Guide fi~r US unreliable results, substantiation is based largr~v Army A.Ircrafi, (U), USAAMRDL TR 71upon the results of structural tests. 41A(U) and Tk 714IB(C), Eustia Directoratc, November 1971. TEnNGVA, 11-82 '. W. Turbow, et al, Crash Suriiwul Design Guide, 114 TS~NG2. Portions of th~e structure will be tested after brief USAAVLABS TR 71-72, US Army Aviation preliminary analysis. These include fatigue-loaded MAI614ati Laboratories, Fort Eustis, VA, October ptimar structure, redundant structures, components 1971 (Revised). manufactured using new processes or materials, and 3. E. F. Bruhn ejt'al, Analysis and Design of Flight castirgs with low safety factors. Vehicle S:ructures, Tri-State Offset Co., CincaIF.nati, 011, 1965. Oth-.r poktions of the structure will require de4. G. E. Maddox, et al, Stress Analysis Manual, tailed analysis and backup testing. Where expensive AFFDL-TR-69-42, Air Forc Systems Cornprocz sscs and/oi lasge, crastly structures are involved, extensive analysis is necessary in order to promand, WPAFH, OH, February 1970. 1-3
within the requiredi limits. In special cam structures that are redundant or subject to fatigue it may be. neceasary to employ both methods. -e~g.,
vAt
JAMCP
)
-.
-
5. J. A. Vichrne awd T. L. Houas 46Huicopwr Cargo Roausin", ;ornal of the American Helicopter Sociny, 10, No. 3,41-47, July 1967. 6. J. A. Vichnees. Integral Helicopter Caego Reagswh Syutemw. USAAVLAr-S TR 69-M6, US Army Aviation Materie. LAboratories, Fort
fJ
11-14
Enatis. VA. 1%69. 7. R. H. MacNeal and C. W. MtrCocmik, Tfe NASTIL4N Ca~~puirr Pn'gWmn for Sigmvwda Andnuysi. SAE Papor No. W6~17Z No.iWoag Ma Aeaw~~adic asid Spacc &Wbwfr4V &w.wkf no*i MwW Los Aapes. CA, Qzco~a 1%0.
CHAPTER 12
°
LANDING GEAR SUBSYSTEM
124 LWST OF SYMBOLS - tire
D
- outsdo diamcier, in.
t W
• 121
GAR
of lakes, the use of some type of water-landing sear is equired. Depending upon the mission requirements,
am, pata in.'
A,
one of the following options could be seece:
dh in.
1. A type with full seaplane capability (either floats or a boat hull) and having little or no land
TPESCapability
Convewtional helicopter landing gear configura-
2. An amphibious type, with either floats or boat hull plus land gear
desired configuration will have been selected during preliminary design. The detail designer's objective is
fabric (inflatable) floats that also are capable of being used for land operations
vide the beat performance for the least weight, cost, and waternance. The system snecification and mis-
packaged floats that can be inflated for water landings if desired. With this option, the helicopter still
specifying the following:
state, it may or may not be possible to take off from
tions inchlue wheel, skid. and float. Normally, the
Stc
verify a type, or combination or types, that pro-
' •
sion requirerments will influence the selection by
1. The environmental and operational landing
cnditions, which will indicate the toughness of
4. A type with secondary water capability. using
maintains its lauid gear. 1Dcpendin% upon the wea
the water.
5. A ditching capability, which keeps the helicop-
cV.Aing tormin
sortd tLhe r•quirments for snow or water landing caability 2. Descent velocities, allowable load factors, and ground ckrance, which together with weight and CG location, will dictate gear location, size, and axle vertical travel 3. Purformance, which may require low-drag or re-
ter upright (although it may be partially submerged) long enough for pcrboroun to evacuatc. Common practice here is to make provisions on the helicopter to attach the flotation gear and then to piovide the major portion as a kit; thus, if the helicopter mission is not over water, only a small weight penalty is incuffed.
trtble landing par, thus requiring a ompac! gear configuration
Skis are required for operation from snow- or icecovered areas. Current practict is to maintain the
4. Overload conditions and growth factors, which
normal landing gear and to adapt the ski to it. This
woid require a gear with increased capacity. As a rule, the skid gear is lighter in weight, is less
gives the helicopter a greater versatility and also improves iki life by preventing the skis from scraping
expanuavv to maA.-
'"
3. A type with primary water capability, using
' a.d
",pIace, a'd req"ire
ku mainweasuc than other types. One disadvantage of the skid gear is the necessity for special groundhandling wheel or dollies for moving the inoperative helicopter. In addition, a method of raising the hilicopter to permit imitallation of these wheels mutt be provided. Some vAicds carry this equipment along and theeb safler the additional weight and drag penalties. Th•whbee ear provides the capability taud directionil control neoemary to maneuver the hldicopte lading and takeoft. on the rod, a well as during In addition, this type of pear allows a running takeoff, which provides additional lift and thus permits an inctease in allowable payload in the event of an overload condition. The •ia am absorb part of the impact energy during Iandig and also car act as a aishice ovw obstade during taxi operations. Where the primnary mission of the helicopter involves flixht o¢er waterways or areas with a number
when encot-sintering arcs bare of snow. Bear paws are
smaller than skis and are used in marsh or bog areas, where normal wheels or skid gear would not be capable of supporting the helicopter weight. 1
WHEEL GEAR
12...1 GC~hl There are four types of wheel gear in present use: 1. The conventions! or tail wheel type 2. The tricycle or nose gear type, with and without tail bumper 3. The quadricycle arrangement 4. The bicycle type gear with outrigger wheels. The recommended location and rollover angles for the tall wheel amid nose gear types are shown in Figs. 12-I and 12-2, respectively. Because of the requirement for Army helicopters to operate on or from surfaceis with as much as 15 deg slope, the turnover angle in any direction should be at least 30 deg (par. 13-
"12-1
AMCP 7~O
41time
4Y
1.1.8, AMCP 706-291). Other than turnover angles, are no specific requirements for or limitations on the location of the individual gear fore and aft of the CG. The quadricycle and bicycle arrangcment3 may fall into either now. or tail category; thus, the same guidelines shown in Figs. 12-1 and 12-2 apply. The tail gear arrangement automatically. wvith no increaw in weight, provides protcction to thc tail rotor during tatil-down landings, whereas the nose gear araimnganet uually requires a tail bumper. or, in esMone, 1Afourth gear. However, no one arrangement
inherently ic superior or preferable. Many designK criteria and op-rational requirements - such as sa&e ty, ground handling, and transportability - must be considered during selection of the landing gear arrangement for a new helicopter. Because the pertinent trade-off studies will have been made and the optimum type of gear selected in the preliminary design phase, the following items should be required at the inception of the detail design phase: I. Gear and location, rollover angles, and other
LINE OF WHEEL MOTION DUE TO SHOCK STRUT TRAVEL
I
SEE NOTE 1
.
LikiNr
_
NOSESSTAE0
11
_
NOTE
i r en
ra AIIAi
IQAN _-a
_
5
'inlO
U2OSHC-TU SEEV
NOTEES
2~
~ THE/MI
LINE OF
LES THALNE OF WHEEL MOIO
Ii~NO
TO0
VEMSTBATN CKTRTT MOTIONDUTOS WHEEL. DUE
AG SHOCK TAE STRUT A
R'
90OS WEIGHT ABO THE CHVE H RORIOTLHE OPITIMUC AL SA UEIS 60`R AD OEOg.LE 35
H TAIL WHEELTRAINL D
122Fpigur 12-1, -Io~ft,2
EAXIE MUST BE GRLIEAT
D THRAN
5t ;,
D ATANAN TH TALWELdoAgTf. OF
WhceW LcAudw sod Modems for Tail Wheel
Hielicapitrs
II
1%
~AMC
--m-lab shown in Figs. 12-1 and 12-2 mum be met.
40
truts and olus in a
6. Bakig cpablity adequate for both stoping and purking th healopter on a reqired ulope mum
Mtatec eltion mot be adequate for the Operational mviroamaa astipated for the upn • hIcopvr
be establ~ished. 7. FIn grunid loads mum be alolated, and the
of 6 in. with all abslom does. A minimum deawancs oai and tre bottomed out. or wit any om Wol sad ti.e bottamd out and all othe in a staic po&onis recomnumdad. 3. TMras"must be adequate to satity pound fio. tadkm requiemeunts at the most adve CO positon. SA dhmmy analysis sbould be performed to in& danmwthat dw Mbut combination oftrail anglead pift is being used. 5. A pound resonnce analysis must be performed to asse that spnog ratse wd damping mosftiits of tke also and the fire have the popervolaum to keep the helid0pt out of the resonant frequm00 range (Sue per. 12-3).
meat atmail loeds on the gmar damrnined.
2. Oocuad dannms with
M1-3.113 CepeeM Dsdp wW Sdaala .1 132-1.LLI Ground fotation requirements. operae% terrain nterim arm and nvirament, Ond gioundrsosva tire selection. Design the main fk*os that influme and construction featur and qualification tests as listed in MIL-T-5041 apply to aircmft tirn. When thee tirs arc use on helicoptemr standardaircraft rating may be a4justed as shown in Table 12-I. Because uf the low landing speeds and short taxiit& didtancr required in helicopter operations, verticel tinr dection gram than for standard aircraft
SEE NOTE I
j ~
ANGLE MUST HIHIS EXCEED 60e
rNOT
SEE
f-SEE NOTE 2
NOT LESS THAN 30° NOTES 1 CG LOCATION TO BE EITHER AT WEIGHT EMFTY OR AT BASIC STRUCTURAL DESIGN GRO3S WEIGHT, WHICHEVER IS MORE CRITICAL, STATIC TYRE AND OLEO DEFLECTIONS AT BAS•C STRUCTURAL DESIGN GROSS WEIGHT ARE TO BE USED FOR ALL CONDITION 3. 2 THE LINE OF WHEEL MOTION DUE TO SHOCK STRUT TRAVEL PAUST BE AT AN ANGLE OF FROM 0 To 7.5deg AFT OF THE VERTICAL 3 REFER TO PAR. 12-2.2.3. FOR TRAIL DISTANCE REOUIREMBENTS. 113.. 12.2.
Whed Lea.0tam a1d Moshe for Nose WIed Hedlcoiters 12-3
AMCP
Wt-
TABLE 12-1. LOAD FACTORS FOR HELICOPTER TIRES
TIRE OUTSIDE DIAMETER
*LOAD FACTOR
*NORMAL INFLATION FACTOR
26 in.AND UNDER
1.67
1.59
OVER 26 in.
1Zo
1.50
*TO BE APPLIED TO BOTH STATIC AN~D DYNAMIC RATINGS
* *specified
is permissible at static load. Type III tires are classified as low-speed, low-pressure tires havins a [Mre cross section, and are used whein. good futst.. lion capabilities are desired. Type Vil tire arm classified as high-pressure, high-speed utirs with at smaller cross section than Type IIl types. Typo VII tires have a high lateral spring rate, which may be required to control ground-resonance tecndincics. In addition, their smaller size allows their us,: in highspeed compound aircraft where limited space within the airframe for gear retraction requires a smaller, more compact tire. For Tyme 11; tires a tire deflte'tion not to exceed 40% is allowed; for Type VI I tires a deflection of 37% is permitted. Tire. efficiency for calculation of gear energy absorption is taken as 45%. To determine tire stroke under load, a tire deflection versus load curve should be used. This curve normally is available from the tire manufacturer for any standard tire size. Ground flotetion is the meaure of the ability of a tire to remain upon the surface without causing breakdown or failure of the soil under the tire. The factors affecting ground flotation arc: l.LoAdUEAUuepe-u 2.Tire pressure 3. Wheel spacing 4. Strength of surface and subsoil. Flotation requireinent.~ fifr a helicopter usually are dependent upon its primaxy miasion. Thic. usually is as a California ['caring Ratio (CBR) value coupled with the numnber of patses the hel1icopter ca make before failure of the surface occues. The method Df analysis or the flotation capabifities of a helicopter on xicprepared surfaces is discussed in Ref. 1; additiona~ll'v it is imrortant to note that the deceleration rate due xo braking should be 6 ft/sec2 fox rotary-winil aircraft. A mcre priwise method for 6eternhing tire contact area ti is A, - 2.0 12-4
WD,inl.'
(12-1)
where W -tire width, in. D X6outside diameter, in. 2.0 - anpirical constant, in. This equation has been derived by analysis of'& curve faired through teat data from current tire. Sufficient clearance must be provided in the "ading gear drsign to p~event tire chafing against the airframe or gear structure under all conditions of. loading and operation. MIL-STD-878 establishes the procedure for determining the clearances required dx.. tn arowtlh nf firms a d inrrepase in diamefr duaw in centrifugal force. In addition to theiot values, a 4% increase .n section width and height shall be allowed to compcnsatr. for the overinflation allowance for tirs used on helicopters. Anl allowance for growth in gross weight (25% minimum) should be made when wheel and tire sizes are selected and 6earwaxos ar establistiod. To provide for such wtight growth. the additfon of plies to incrrase the load rating of a tire otherwise suitable for the design and/or dynamic lods is acceptable. A change of wheel and/or tire sizes during the service life of the helicopter can be orqarc C103sp-toW~r
AA3
Nowir A---
airframe structural components and therefore should bc avoided. Wheel. 12-1.1.. Factors that affect whorl design include notch smasitivity. fatigue, corrosion, hecat damage due to sestained braiking, and wheel disintegration due to tire failure. Current practice is to xpecify dhe use of forged aluminum alloy materials (usually 2014-T6) for new designs. Cast magnesium is not desirable because of its susceptib~lity to corrosion and its tindency to fragment or shatter upon fail~ire. R-elicoplte wheels shall be designed and xesled to the requirements of MIL.W-5013. The tated load for the wheel shall be equal to or greater than the maximum static load to which the wheel will be subjected at rcxaxiinum towing or alternate grots weight, which ever is greter. Roll tests shall be for 250 mi woin-
Pared to 1500 mi for fixed-wins aircraft Bearing sime can be aestbighed by the mthods cullined in Ref, 2 or a comparable roller bearin havdbook. A seall or other means of Preventing water from entering the whee beAring must, be providw to prevet bearing corrosion. M1.143i~ Sube sturat The most commonly-used devices for absorbing eaneg during a landing. and for supporting the hellcopter during taxiing and ground operations, are me. dusncc springs, liquid springs, and air-oil struts. Other methoidsatc absorbing energy (usually referred to Ik-secondary energy or onetime application usage) 41= i. Crusbaltle structures, e.g., honeycomb 2. Fracion deVices i.e., two Materials in cotc with escb otheiryeofihc Sdevices e0g., meWa drawn through an 3. W okfor undrsad 4. Cutting c viaes, e~g., a shcrp e*dgesicing throusih mewa as it moves 1._ .*. &am 4-', I& .. cornaio, i. wow"% %W%
*wh110
beyond the elastic limit. I-)Thes deviot arn mentioned as possible solu-tions An energ absorption beyond the normal land-
mug Par cauaility
is deid The meet commonly-used mechanical spring is th
such a a "4ecuper atrunit must be provided to limit the PrOSSUM file d"n to high temperature. A liquid spring is perticuiarly Well sated for short stroke applications with levee gear affrragmets, but any length sttabe can be achieved by jproptr comnbinatiosa of piston rod diameter snd total volume of oil. In fact, with tail rod liquid spiungs, exceptionally lons strokes can be achieved. For servicing a liquid spring without jacking the helicopter a charging pressure of 20.000 psi is required, while the maximu chatgi" pressure required probably is 2=0 pei when the sp~ring unit is unloaded and fully extended by jacking the helicopter. In the air-oil strmt a chamber of compressed air is used to restore the strut to a static position and to provide a cushion while taxiing and ma acuvering the helicopter on the ground (Fig. 123.Teefcec of bore sson ydo et to be between 80 atii 93%. A conservatively low value tIA shock absorber efficiency should be us.4d together witl he specified reserve energy descet ve6 locity and the design limit landing loal factor, to es-
tablish tie maximum strut stroke required to absorb th toud roVt eftensy. Thu d r c~a'30ik-90-likdrisdii-
cussed in pri. 4-10.3, AMCP 706-201.L The lower values of efficiency are applicable to
struts containing fie orifices.1The higher values amc
achieved by the introduction of a metering pin. This configuration providca a variable oririce, which arm cantleve type, with either a flat or a tubular cross "sction. Thee can be considered to have a firouar load METERING PIN deflection cuarve if they are not stressed boyor.d fth elatic: Limi. Gther types of mechanical springs -ORIFICE AND BEARING
gaab as Belleville washers, ring sprngs, and rubber -have
nonlinear load deflection d~aramn. D~am-
ping is small (usually due only to lees of energy at the
suport, or frictional fores between dlemenits). Because the ipring return far more energy than they
l~II.-7~
SNUBBER
0
PSO
dissipate the use of spring landng gear usually. 0 limited to ligM aircrft. The lack of damping alao incraes heliklioodofground resonaace. Both the liquid spring and the air-oil strut absorb eNWWg by the dnshpot principle, forcing fluid underLOEBARN presur through an orifice. In either application thOEReARN a&M of the orifice may be fixed or may vary with dis-
ISO
CIsINE
placement, as a metering pin of varying diameterI
*
moves through a fixed orifice. lnathe; liquid Spring the compmasibilityof dAe fluid is vted to stoM energ. Special wonsderation must be glv- Ain the gland desigm to the high pressure at whichi these units operate. The sensitivity to changes in temperature also mumt be considered. For example to accommo01date low temperatures the unit prob&Maby will be prumemrined in the extended. or unlodecqdtion Wo 20 psi and s%"Wia feature
YRUI
I.
e_
LI
SEPARATOR PISTON AIR
Ilv. 12-3.
Typical M.4N1 Sermi I Static Feellie 12-5
k
AMCP 9-0 be large at the point at which spin-up, or ýaigh drag, loadsI ame applied. The orifice area then decreases frtherw on in fthstroke and more energy isaboorbod at lower load levels. However, this type of strut is more complex and the implications of increased weight and cost relative to a strut with fixed orifice must be weighied. The desigiiA of air-oil struts Aai,) be in accordance with MIL.L-8552. with two excentions: 1. A seal other thar. the standard 0-ring to avoid spiral failure 2. A specal sraiper ring that is morce 6lToctive in keeping out dirt than the specified MS 33675 scraper rinip. The latter have no means of preventing dirt from entering thertrut paut the outsiei of the scraper, which fits loo~elY. The first, step in the design of the oleo strut is to determine the size of the piston. The ilatic load and santic presure define the %r^a and comelunl8~IA~y th diameter. of tLA piston rod. The static pressure is determined from an isothermal air compressionstroke cuive drveloped for the ltrut. A press5ur of 3000 psi can be asdumed at the cornpreutsed position. with provision in the strut desg* ior a maximum of 4000 pai to allow for growth tincreased groas weight) of tha helicopter. A compi usson. ratio of 2.5-3.0;: 1.0 from the static to tlhe compressed position commonly is applied during preliminary design, with the vallue sclected being dependent upon the landing load factor. Thus, the static air pressure used in the deterraination of piston diamcter will be in the range of 1000 to 12WO psi. To avoid full exlension of the gear when the halii copter is lightly loadtd, the pressure in the extended
[
G;OATHLi MAXIMUMJ ST1%J1
3000
L
position is aln at approximately 25% of the static pressure. The static position is sat between 66 and 65% of the strut etroke, with 83% providing snoother taxiing. Am an example. for a strut with a total stroke of 12 in., the static position would be 10 in. (03% of 12 in.) from the fully extendod position. A typical shock struct compression curve is shown in Fig. 12-4. When the static positions of the individual struts are being selected, the overall landing gear configuration also must be examined to assure that all applicable groundl clearanc requirements, including flat tire and fla. stmt conditions, are met. As prvously noted, the piston area is determined by the static load and the static pressure. The wall thickness of the piston rod is determinied by the cnitical combination of bending moment and compres,sion load. Of the ground load%bpccflod (pars. 4-5 and4-, AMCP 706-201) the braking conditions or obstruction load condition probably will produce the maxiimw' value of bending moment in the piston rod, whic~h occurs at thc po~int that the piaton enters the lower bearzia. Materials with ultimate tensile str-ongth of 260,000 psi or greater should be considered for *h. piston rod, subject to the approval of the procuring activity (MILL-L552). The ratio of piston rod diameter to wall diicknecs also musi ba examined to assure that the design is nather unnecessrily heavy nor impractically difficult (hence expensive) to manul'acture. Should the wall thickness as determined by the critical loading condition, prodluce an unacceptably hecavy strut, the conditions governing the static position should be re-cvaluated. There are two ways in which a shock normally is mounted. cantilever anad univtersa (Fig. 12-5). Tht,
POSITION
NORSAL AIR CURVE
2000
SAI
1000 0 0
25
POSITION
50
66
83
'00
STU TRK,%CANTILEVER GEAR
F~gur 12-4. Typical lsmissisCr Shock Stmt Air Campcaesulsi Cur' 12-6
ARTICULATING GEAR WITH UNIVERSAL -MOUNTED STRUT
Figrm MS-. C~ausklew MWd UNIrermlSei
.
esaghiee deep is used mom, ommonly because Itslih.I& loweight ame tbwonl lar beanding moments imy be Induce under leImft condltim that inGOWOO~eait5 Odhe than al104g thm Osideords elud" anis of the strut. ThWm beaft skmoments an ere the stand4point Of the r~uoltt 3Naa Vot o0* 111om1 stoma,. but also becamse they increase tie Possibility of likgek through the strut bearing and increse strut brelak. Tuhis *icon can cmus erratic strut operastie, durNg landn and taiing operations. Howe ier~ of pm~ebusering Uiakiu mrome locatioin of Saa Intan bmearing ar low busring pressures an@ sIdas Wantt~iV& For caatilsver-maountmid Amtrut, the bearings should bespaedl so that bartingi Sart don not exceed 6=I~ under the du~p "idload. Ma distanc between the outermost ends of the upper and lmwe benflfl with the strut Mly extended should be at least 2.75 time the piston diamseter. For universal Wpi ne) Osp struts, which have littl5tfne~lh Wd yt ~a bdm teace is only I.2 tMdes the pistoo dkmeer. With thus paramesters known, the length of the st" can be dawemined. The length ofa catier strtw Am thn do extended p~on sioa1 equal to (Fig. 124):
-s
)
1corwec
7~~ In. APPRO)XIMATELY
STROK(E
II
I I
2.75 x (piston diameter) + 2x (sdroae) + (allowan= for upper cap and low aske cmn inhown in Fig. 1246 appinuiMia*d This Aa c~ws, 4.0 in.. will vary with strut sin. For univernalmone stus th ent iscalculatee in the Soots mainer as for the cantilever strut. excet that the lowervalue of minimum bear*n Speclag Is use. The elimination of bending loads by a levered suspenumon thus reults in a shorteF stMut The0 meSOf this 00011ig1uratlon is reamnasndud for tail Par whats losng Mrai arms san esirable to reduce shimmy tendenices. A snubber or reboun ring is provided to reduce th imac fojeuo the par during Sudden exter lom, Such as those encountered on takeoff and landln reoud The strut in the fuly extended position ead not oUwchb the ground must be Capble Of reac. ting to a rebound load factor of 20) times the unrung weigit Of the landing gear, or three times the lead Induced'by air pressure within the gear, whichawL greater. M~r of air and oil lb the olso can camn loss of gttafliec as the hydraulic fluid becomes awasnaL AbmL foasmles of the fluid takes nlnes durnga esesing the possibility of having dwn in. servkkinu. amount of fluld In the ollso end thus reducing Some of the eaergy-abeorptlon capabilit7 or'the samt. Current prac'.ce isto Insall aseparator piston, which Gute two separate chambers; thus, no mixing of air and oil is pceable. The housing or outer cylinder is usually oi~ forged alumintem construction and consists of bosse or lugs to provide the method of attachment to the airframe. A sat of torque arms is necessary to reactw the tmique lowoo on the gear dimt to pivoting. ~~and to maiattaln wheel alignment by joining the . A..f .. n I Ln .w..
EXMDDain
2.75 XPISTON DIA FOR SM4T CNIEE TU LENTH1.25 XPISTON DIA FOR UNIVERSAL. STRUJT
STOK
\ 'C)
-.
Npr.12.
STROKE
2.0 in. APPROXIMATELY
Dsesrsahdead
of Strut 9troke
w
r . ......
axial doom. of freedom. Main gear torque NrmS uwally an cosrce or forged aluminum, with the angl between the arms limited to 135 dog. Nose gear torque arms, being Susceptible to shimmy, may be of ste construction with a hew angle, and may use a knuckle deApg (multiple hats) at tho apex end. Trhe dusig of nose and tail wheels musm include centering spriupg to insure proper alignment of the wheo* during lending and sufficient damping of angular motion to prevent shimmy during mn-on landing and zac14g. Shimmy is a self-excited, rapid oscrillation of the leding gear that occiurs at Or above critical landing spodxs. Basically, shimmy is the result of a lateral mimaianmen' between the helicopter CG and the ceuter of contact of the tire with the ground. The gea isdef kte to one Side, but a restoring forme ~due to the elasticity of the gea and its supposting.
structure causes tIte wheel to move back. It then over-
13.7
-
Af.
AMGP70W
shoots the oenter, with a subsequent lateral misalignmoet ca the other side, An analysis of shimmy is quite complicated if an attempt is made to incorporate a11 of the variables; however. a aimplified approach that calculates the speed at which shimmy starts to occur has proven acceptable for landing speeids below 100 ki. This method uses Moreland's stability critria (Rd.~ 3), which express the eqiusions of shimmy in a nondimensional form contaning all of the major aircraft parameters. A digital computer program has bee developed to indicate the stable and unstable regions for any given aircraft configuration. To reduce the possibility of shimmy, several design practices based upon experimental work should be obseerved for helicopters equipped with nose geaup: 1. The trail distance (Fig. 12-2) should b,.- less thin 8%oir greater than 50% of the tire diameter. 2. Dampers mounted at the wheels or at the strut are acceptable. For dual wheels, a damper connecting the two wheels is preferred because the amount of play in the syttem thereby isminimized. Dual wheels are preferred for dynamic loads above 20,000 lb. 3. A shcort trail distance usually requires more deeiethen a
Ie., *~I
ar &`i------
Imting oc acmiarticulsting Sear with a long trail arm is
probanly lighter than a gear with a short trail arm. 4. Hydraulic viscous damping is preferred. Friction damping is not desirable duc to the large vaniations, encountered, 5. Torqt arms should be ui stiff as possible. The apex should be a knuckle design to avoid any offset in the line of action of the two members. The use of steel instead of aluminum should be considered, Insiglerotr te csef cofigratonsthetan rotor provides excellent c',ntrol o( the helicopter duiv&ini isndiiidi-, uakeorfM, and iaxiing; therefore, steering of the aO~or tail wheel is not necessary. A
device should be providrA, controllable by the pilot,
to lock the tail wheel in a trail positucn during landmaS and takeoff to assist in directional control. No gea helicopters are inherently stable and, therefore, usually require no lock. Both nose and tail gears should cortain cams or other centering devices to maintain the gear in a trail position prior to landing. As a viuue, antiskid devices are not required for vehicles with landing speeds below 100 kt. 12-1,12.4 Brakes Braking system requirements are governed by the system specification, and by MIL-fl-8584 anid MILW-501 3. Current cargo and crane helicopters wec Type IV systems because the energy-absorption requirtment~s dictate a power-operated system. This 12-8
type of system must operate from either sot of pedals. and must perink parking of a lO-deg slope without application of external power. Main pear wheel braking usually its sufficient to achieve the specified deceleration; howeve, each hell copter must be analyzed indvidually to determine brake adequacy. A parking brake handle accessible to both pilot and copilot must actuatc a parking biake with pressure sufficient to hold the helicopter on the speOcified slcpe with the power off. Sizing of brakes is dictated by prior constraints in additicni to the applicabkc Military Specification. The size of the wheel and tire. previoul~y selected, defines the volume available for the brake assembly, and hence the area of the friction pads and the number of actuating cylinders. Master brake cylinder sizing and detail designt are dependent upon the brake pressure and actuation volume requirement, thc speed of response, the linkage ratio to the pedals, and the available pressure supply. 12-1.2 SKID GEAR 22A
Ger-
The major advantages of skid gear are lightweight, low cost, and simplicity.. The initial cost of skid gear is leew than that of a cvrnventional oleo gear. The~g elimination of wheels, tires, brakes, and braking system oleo results in reduced maintenaicr. The dieadvantages arc the need for support wheels or dollits to handle the helicoptea on the ground, a limited running-landing capability, the inability to perform running takeoffs and to thereby inrreace payload with the resultant increased lift, and the high rate of wear of the bottom of the skids. Skid gear is ustd on many lightweight helicopters where the normal landing energy ii stored in tubuler or rectangular spring members. For harder landings. the landing energy is akiorbcd by perumtnent deformation of the spring members. The- static deflections of' the skid gear usu~al!') arc ls hntoco losa.Teercec fsi gear is approximately 50% until the load in the i~pring member exceeds the elastic limit. When the load is abethyilsrnthotemmeteef. ciency of the skid gear is comparable to that of the conventional gear. Ground clearances of structure, control surfaces, or external items for skid gpar shall noi:x~ less than 6 in. witi skids flat on the ground in a static position. Wheels 12-1.1.2 Groud..bsadiI Location and number of whiels are dependent
-
)
AMCP 7064202
uopn helicopter weight and CO location. Units shcu4~ have the capability of being removed emsil and quickly. Whoek mhould be capable of rolling over a 4-in, elistacle. A braking device is not required. M124±3 Scuff Plane Facors that affect skid wear are speed of landing, beaing press, *re. rni~Aanoc of skid material to abrasion, and type of kendin surface Removable wear plates r-hould be located at critical wear points along the bottom of the skid to prevent permanent darmagc to the skid and zupporting structure when a laading is made on a hard surface at a speed of 35 kt. These wear clates must be designed to prevent excessive digging into the pavement marfaqc- as a result ur the scuffing acion rerulting from gear motion. Steel wear strips semto give thn best ruAuhs. 12-1.
9
RETkACrABLE GEAR G61011gear 12-1-l requirements for th retraction and Extension kusd.. ac-nunbnn syseeq arm viven by the helicopter
system, or detail specification.
The gear x.tuzation system must include a mechanical lock ateither extreme of travel; must provide an indication of gear status - i.e.. up, down, or in transit - and mumt provide a method for emergency akctuaion Reracioncanbeaccomplished by folding tuwatin aetrlaction canly be fowad, ateall, ftor orby elecopng he ear along a fixed oleo axis. iForwa d retraction is favored for the main gear in order to pernall the airstream to assist in anergecy extension. The telescopic retraction method may be wmployed on the nose gear, MUM..LAP%.&UW~1 rtu
Pnose
The design of the actuating systemls dWU be in sccordmnce with MIL-C-5503 and MIL-H4775. However, the specified seails (MS 25771) and scraper rings (MS 28776) should be replaced where possible. 0ring shaft seals are prone to spiral failure because the seal works on an unlubricatod shaft. Improved &arvim~ life has bcqr achieved by using a sead consisting of a T-shaped elastomner supported by two Teflon backup rings. MS )crapor allow entrance of sand and dirt because the sealing surface is discontinuous. A filled Teflon acraper. preloaded with an 0-ring, provides longer seal life. 121-3.2 Actuadom Indication usually is provided for three modes on each gear. Engagement of the mechanital up and down locks a~nuaies fth gear up and down indications, respectively, in the cockpit. Disengagement of both up and down locks indicates an in-transit condition by illumin~ating the control handle front within and/or by uncovering a striped "barber pole" indicator. .,
Emergeacy extension of the Sear should be manual. An air bottle may provide the energy necessary ta~ assit in lowering the gear and to overComte air loads on the gear door. Provisions sM/I bea made for emergency extension in the event of loss of hesoiSteSa ydraulic pressure or of failure of the landing gear directional contiol valve. 12-1.4 SKIS AND DEAR PAWS M21.4.1
large, open whWe well in the nose, or because the gea strut length must be controllable in order to tilt the fuselage for cargo loading, A typical retraction system, wherein the VA.- pivots up into a wheel well, contains the following: I. An actuating cylinder on each gear, which either acts bs a drag stmut or drives the drag strut linkage, which in turn actuatas the gear 2. Mechanical up and down locksV 3 Mechanical lock limit switches 4. limit switches on eah gea.- scissors, or on the oleo itself, which are deactivated when the helicopter weight is on the gear. The down IGca limit switch, the landing gear congtrol handle, the scissors sw-tches, and the hydraulic control valve up coil am wired in series so that raising can comnmence only with gear unload. Similarily. the Sup lock limnit switch and the control valve down coil ar wre i sris through the control handi.
General
Skis and bear paws are similar. Bear paws are used
*'-ea.
primarily on snow or soft terrain for nearly vertical descents. Skis are used primarily for landings with some forward speed in snow-covered areas. Skis are VI
7E 9
G4RSATCAD0ALO FRSAI N ALO NOE GEAR DYNAMIC LOAD
3I
200 1 bcI 1001
NOw 1
51AI
AL-EA
OD
160 2000 Iwo 00 120M5 GA OD Flgarn 12-7. Lmudag Gear Static Load vs Ski Drl. Premere M
larger than bear paws, and have a lower bearing pressure, a longer nose section, and a greater lengthto-width ratio. Landing gear static load versus acceptable ski-bearing prosutre data ar shown in Fig. 12. 7. Nvt.t that the main gear has a higher allowable bearing pressure than do the nose or tail gear. The tendency of pilots is to feel for the ground, or land ~tail down, when operating tail wheel hlidcolpter. This increases the effective tail ski load. As a result, relative to the main gear static load, a lower bearing pressure is required to stay above the snow surface. The nose gear must react the dynamic. loads caused by drag and friction forces at the main gear, and thus a lower bearing prosaure is necessary to keep the nose ski from submerging in the snow. A comparison~ of typical missions indicates that lightc. vehicles require a greater mobility than does a heavy cargo helicooter. This is evidenced by the lower bearing pressures found to be acceptable for helicopters with lower gear static loads. Limited available experience indicates that above a bearing pressure of approximately ?50 psf-, ski --
...
L
t
.,.;
-
aj.,.he
bearing pressure for bear paws for soft-woil operation should not exceed 1500 pet (approximataly 10 ps!). The length-to-widtb ratio of skis is not of great importance at the low forward speeds encountered in helicopter operations. A ratio of 2.5:1 c ,rrently is in use for helivopters. On snow-covered terrain, a run4ERN
/
PRESSURE,
35
OF.
lb/:W
ning landing speed of at least 15 kt is required to enable the pilot to maintain a clear field of vision by keeping the helicopter in front of the blizzard createdI by the rotor downwash. Ski friction is due to compacting of the snow (which can vary greatly with the moisture content or density of the snow) and sliding 'riction. A coeffcient of friction of 0.25 hall apply for landing coaditlons and a coefficient of 0.40 &W1lapply for ground-handling conditions (NIIL-A-M62). The relation between snow depth and ski track depth is presented in Fig. 12-8. 12-1.4.2 Imstalladua Current practice is to adapt skis to the standard wheel arrangement. The tire is allowed to protrude through an opening in the ski in order to permit landing on hard-packed snow and it-e. This permits the tire to absorb some of the landing enegy that otherwise would be transmitted through the ski. This design also pernits ground maneuvering of the hellcopter to work arma that normally Are cleared of soft snow. .I .L -- allim. .r sv.uj gmt nr ihawWist Mnntme.tion with a honeycomb or balsa wood core. A long planing nose (Fig. 12-9) keeps the ski from diggng into the snow during landings involving foward speed. The aft portion asoe is raised upward. although~ to a lesser degive, to permit rearward movemeat over snow or other obstacles without digging in or snagging. Replaceable chafing strips may be at. tahdW to the. bottom of the ski to protect it Guom
being scraped arnd damaged if it contacts the ground
while landing or taxiing on snow-dleared surfaces. pedestal rat~ings thtpstintsi
-0
//Z u~2 0
/ W''OO lb/ft2
0
553 hi-laKUU
________________RESTRAINING
CABLE
0 0
12 SKI TRACK DEPTH. ft
3
1~no 124. bmw D"M we nai rakDepth 12-10
U
M
wheel and permit pivoting of the ski above or below its normal horizontal position, thus allowing it to follow the terrain without imposing high loads on the ski or gear structure. Fore and aft cables with adjusters pusation the ski in flight to minimize drag at cruising speed. Cables are attached above the sho-k strut and become slack as the oleo coutpresses upon
BEARING CONTACT LENGTH Figre 11-9. Ski Cenfiguratla
landing, allowing the ski freedom to suck its own position on the mnow. A spring or bunqme cord sanetimes is used to keep the, ski clear of the ground on sow-free runways, alhbough some designs favor a caste wheel at the aft end. 1242 LANDING LOAD ANALYSIS The operational envl'anwaet to which Army helicpesare exposed is sufficiently severe. that it has naese te snki# seedrebowounnenuqto quimremnets spicikied AinMIL-S4696. The applicable deepg cr,ýeuia ame given in Chapter 4, AMCP 706. 20.The requirements fot both symmetrical and "aymmetrical lending conditions are discussed thervin, iresmuch as a comnprehevisive design anely4s of the landing gear subsystem is required uurinS the preliminary design of a new model of helcoter. Landing loads tA0l be determined by a rational analytca procedure that has bean approved by the precuring activity. One such procedure is outlined in
MIL-A4082. The adequacy of the landing gear sub
j
sequently will be demonstrated by drop teat. Th ~~~qualification requirements, including the drop tut,
4
:
2W..
..
Inad tio o the landing loads, the loaui created by wtuiing over obstructions, turning, braking, to'~ft and backing also must be determined using the design caiuria orChapter 4.AMQP706-2D 1.The hendin end ground-handling loads then are distribunted as shea loads, and bending aid torsion moment at selected points on the helicopter. The omens lods nd rsulingfromfliht eds shee &hea omens lods ndrsulingfromRigt leds are distribiuted similarly, and the critical loads at cecti point are determined. ?Worwall-, the maprg-absorbing capability oi the t.
..
~
,
natural frequency in the rolLltrl ado pitching modes. Instability cart result in oecilfktionpM which buildup dufflciently to dkar')y the helicopter in a matter of seconds. Because the theory is uiderstood (Refs. 5 and 6), the problem can be approaclid analyticelky during iCetaiI design. The feiiors governing stability en a combination or rotor and landinig tear parameters. Several of the pa aneters amre ixed by basic consraints beyond the contrCi of the landing Scar designer. The critical pantfieta i.lae rsofenmo fiai .Baewsofe.admmn blades 2. Number of 3. Blade damping about the lag hinge 4. Rotor rpm 5. Fuel mess 6. Fuselage polar mess moment of imnria about the lateral and longitudinal taxe ergoer 7 8. Tire vertical and lateral -kpring rate 9. Oleo strut damping 10. Structural sping rate of gear. ffrtema fciiiyi tme,7 8 ,ad1 0lvofferon themstofexiiliy i 9 aItemsing7 9 and The problem is furthmr complicated by the folwn -ators: I.oniearaitonfthtrendlwpig golosprwinghto byvargsiain othetire and rate Nonuinea rtscaue ycagsi eicpe rs egto 2.TeprntgofVswihtupredbte 2.otheprotgrfgoswegtspotdb h ror 3. The effect of inipri-per servicing on the spring rates of tbe struts and tires 4. Th effect of a flat or soft tire on the spriqg rate. Adti einapoc o sesn tblt n Adti einapoc o sesn tblt n dirtcriining the parameter changes rcouired to
MrUMnd-l iig loadso toU"ticlolo the lanimn-and. ing paer attachment and support points. Should the. sapeified lending or ground-handling loads exceed the flight loads, it usually is appropriate to reil th landing 9wa isnergy-absorbing system to reduce the load factor at the CO. and/or the local loads. The airframe is overweight and structurally inefficient if the lending loaids are critiewi. As with other chairacteristics, it is necessary to examine the trade-off et~ween energy-absorbing-systemn weight and airframe stut-d*ih.lengths. strutura u~.ght.CBR, OFGOUNDspring W AVODANC
)
resolve the problem includes the followin~g prc'-M 1D erterieteltea.rl,:d ort-i .Dtrieheaerlolndonidna rigid-body natural fircquenzies of the airframe, considering Items 5 through 10 plus the aforcni.ntiontd nonlincarities. Operational considerations will deterietelneo'aalbeprmtesi oecs~ A range of oleo spring rates is est -.1i)ii!cd by the dccele&raion loads imposed apoez the airframe anid by the ratio amiong static, wextened, and compressed Tirv ý,tlerction is dictated by terrain, load, and leaving some latitude in tire vertical and lateral rates. The ranoo of oleo-strut damping is esen--rgy-abA OINANCE O GR NDtablished by the oczessity for art efficient natuial RES frecurve. The possible rigid-body RESOANCEsorpticn quencies, and the extremes of the ratio of rotor rumn Ground resonance oceurs due to coupling between to rigid-body natural frequency, should be tabulated. the main totor blades oscillating about their lag 2. A ground resonano; stability plot should be hinges sn.J the airframne excited at its rigid-body 12-11
AA---
I
AM~P 70&M20 janerated (see Ref. 4 and Chapter 5, AMCP 706201). This curve describes a center of instability by dsfininj, as.a!unction of blade properties, the ratio of rotor rpm to undamp~ed airframe rigid-body natural frqeqn.cy that leads to instability. A kbind of instoihity on either skie of the center of instbility is auperimjiosd uwing Colemnan's technique (Ref. 4). The-analyiical tt-lihmque estabhshes instabilixy in the absancc of also- or rotor-blade damping. A cc.-parison with the mau'th of Step I detrmiines what conditions of rotor rpm and strutitand tire sWTiheuucs lead to irn~taoiity. 'Y.Stabilize the*bystemn t,' reducing the width of the inst..bility band to xzio through the introduc~tion of both rotor hub damping and olec damping. Dleutsch's criteria (Ref. 5) specify the pioduct of bladt and o'eo damping riecetusary to ivduce the unstable range to zero. Required strut damping is obtained from Eq. 6, Ref. 5.Strut damping will be non. linear with reaect to Atroke an~d soad if h tapered ractering pin is used, xesulting in a variable orifice Hcem, the damping xI culd be defined as the tangent to the force/velocity zurve at each discrete comn-
__area.
hiotn
.4
j
KA 4,
ef
ohiit
Inadn
T- h-
nn.---.Ifrut" --
-
-.
2. JEEf=c of partially empty fuel tanks 3. Placmnent of passengers, crew, ant carjo 4. Rotor blade lead-lag effect. For hclicopters with a vatcr-takeoff capability, a minimum dlearance of 6 in. between the rot"r bWad"s and the wiater must te provided at the required us state with rotor shut down. This predludes possible damage to the blbde while the helicopter is rolling and drifting during engine kal'down. Sea state is a condition that comprime height of waves, wind velocity, and wave length. Current psuctice in to use significant waveheight valuies, whbh are considered conservative, for &kSin. A graph of wave height versus wind velocity for different se states is shown in Fig. 12-10. When cak-ubating the buoyancy of a float, it is customary to consider the weight of firesh water as 62.4 lb/ft3 . Inverted V-type hull sh-pee ame preferre over flat or round shapes. Inverted V bulls have low water resistance and reasonable aerodynamnic chtaracteristics, whereas floats with circular bottom wtions tend to stick and to exhibit undesirable sproy characterisi
lkfah-ý
vp
CVa)..U
uoundrssonnceDust be avoided (see par. 5-335, AMCP 706-201) includc one blade dsmper inoperative and the combination, on~ a single stirut, of flat tires) and flat strut (shock atbsorbe pressure at zero). These cases are distussed in Ref. 3. The methods of ansplysis aiso are discussed in Chaptee 5. AMCP 706-201, while substantiation requirements are given in AMCP 706-203.
pi~ot's vision or damagng the blades during landing, takeooff, or taxiing should be incorporated in the in-( itial deeiga stage. Spray deflectors arc a possible solutioji, but they result in increased drug, i.e., poorer performance. Buoyancy of float hulls i3 dictated primarily by interior cabin size, with the result that hull helicopters generally have substantial amounts of !-xcms buoyancy.
12-4
12-4.2 PRIECPBLT
WATER-IANDINC CAPABILITY
12-4.1 GENERAL The design criteria applicable to water landings tire a sink speed of 8 fpo in ~ccmoination with 2/3 rotor lift, and nppropiiate hzad moment and diag at the basic structural design gross weight. Specific landing coniditions to which these criteria apply are: 1. Zero forward speed 2. Forward speed of 30 kt 3 Asymmetrical diop, with the huhi rollee~ 10 deg and no forward speed ~~4. A forwp'1Ispeed of 30 ktand ayaw anglo of 15 dieg 5. A forward speed of 30 kt and w~se-up pitch angles of 3, 6, hkd9 4!tg W!.en detenb..ing the lateral stability of the helicopter, the following latecal imbalances must be con. sidered: 1. Lateral displacement of the helicopter CG froui the centerline as inherent in the construction of the helicoptcr 12-12
For water~b-bu helicoptmr, the basic fletation dasign will consist or metal flo~ats; or bull-shaped fuse. lage witl, some form of owitiggers for stability. The hull and auxiliary float must h- ye enough water. tight comrpartments so that, with any single compartment of the hail or float flooded, the buoyancy of the heclicopter still will provide sufakient stability to prev-rnt capoizing in theseamstate in -'hich it is toopeivite. The high CG inherent in helicoipter design. together with the large droop of the blades, minket, helicoptar operation in the open ocean sea state difficult without the imposition of large performance and weight Venalties to obtajii the requited stabiltiy and blade cea&-ance (Fig. 12-10). The use of a sanchor to maintain. a heading into the wind and waves is one approach to improving the roll azability for cpcr.-&n in OiB sea state. 12-4.3 ADDITONAL CAPABILIT Many helicopters are primarily land-based, with
V.
I
)
AMCP 70&202
some phase of their operation performed on sheltered or inland waterways. For this type of operation, hull and auxiliary floats, if used, must be divided into compartments in such a manner that, if any one ccmpertment is .1ooded, the buoyancy of the heliv..ptc' still will provide sufficient stability to prevent capsizing It floats are used, the buoyancy noceasary to anpport the maximum weight of the helico'pter in fresh water must be exceeded by 50% for &ingirfloats and 60% for dual floats, The most straightforward appr( ich to an junphibious helicopter is to incorporate a watertight hull with more than adequate buoyancy and with little Performance orweih: penalty. Auxiliary floats must provide lateral stability for the sea state condifion specfied. The landing gear can be retracted into the swum ~ 7 floats to reduce drag. Another %&y to obtain a water capability for land. based holicopters is to add separate fabric (bag) or metal floats. The wheels and shack absorbers are at tached to the floats in order to keep weight to a mini-
F~ 1 Ti 12-4.4 EMERGENCY FLOTATION ICAPABIILFTy
12
I-
I
8
I-~.*4
~The purpose of emergency floaw is to enable the
I
OPEN OCEAN
5.
helicopter to, rentain afloa~t long enough for the occucapable ofremaining afloat in a condition of sea-state three if the normal helicopter mission is flown over sheltered area or inland waterways; or in a condition of fPvc if missionr arm parforined over open waters. ~$~\ Two classes of floats have emneigenry flotation ~~X\Q capability: inflatable floatz vrid, 4i0. hrin floats. The inflatable floats, which cm.' be -to4-t?di ':A~ded at the .~
~\\
I
"7'),,
I-~~-~-.--.
f.
3 31
2
,'foi-
SHELTERED AREA/ SEA STATE I
Ir
mum. A mingle main float is not usid on a helicopter dtue to the absence of a wing or othoi' striecture to support the auxiliary floats that are necessary for lateral stability. The permanently inflated bag floats art attached to the airframe by tubular members that s6ipport two fore-and-aft tubes to which the floats are attacked. Thesw latter tubes Vzovide both stiffacss and stability for the floats. The undlerside of the float may have an addiuional tubular member to accommodate wear. This also can serve as an attaching point for a damping unit if such a device is requirod to damp out any tep iency of the helicopter to shake or bounce. Larger amphibious helicopterb have both floats and wheels. The wheels keep the float clear of the ground during land operations in order to eliminate chafing and wear on the under side of the float an6t. o allow taxiing to loading areas without thz need for doilies or other speia: ground-handling equipment. Floats, either primary or auxiliary, must be able to withstand the maximiun pressure diffe-rontial that might be developed at the maximum flight attitude without exceeding limit pressure.
3ý
larking ga.rit
asaft%,~Ca.,
-'*-
Gf o.-
th..
inflatior.. Inflatable fliuts ut'ý a fig' t and compac' system, with a drag increase tidal is only ak fiaction of that of permanent floats. Inflainble floats cnmaintain the fslaecerothwtrathat, in the event of a forged landing in relatively calm waters, repairs of a minor nature may be made to allow the helicopter to take off again and complete its i4.. ssion. Be4cause the floats are inflated onaly a few
hundred feet above the water (inflation time ranges
from 3 to 7 sac), pressure change with altitude is not a
I
~design farter Floats normally are pressurized at _______from
0OLo 0 FlIM* 12-1.
1,
l 10iJ
______
30 20UCTYk 30
DelWmon of F'es State Coudlions
0.75 to ;..25 psi. For use with skid pear, these floats are cylindrical shaped tind are stowed on top of each skid. For wheel gear, the float is doughniashaped and stored in a metal cage awot~nd each wheel. -VI'loats usually permit the water level to cover tht: floor of the crew or passenger compartmciii. This his the effxts of: 12-13
F
1 Lwar o.
the CO, theeby improving lateral
pitchingan be oesrved. Methods o(fIindu timSNJ
stability rotor subela'e or ohw ule dharactuls"2.Re•ikt g thebweilght of the supberged portion tica, if they exit, also can bc evaluated. of the helicopter by the buoyant force of the dispkoed water, thereby requiring lem buoyancs fromntRE lgCTS the flotation bags 3. Keeping the helioptr intacat t the crew 1.D. H. Gray and D. E. Wflaisamk Evhd•am qrA/, can be Frmed and a salae operation can be undercrq4 L.UdJJ"W c G"W, mWfd Fleem a taken. tice for Opemritk from Unuwfamd So0 Airf . Nonmally, two main ditching floats are forwrd of SEFL Repot No. 167, Air Form System Coalthe CO and are powered by a cool-lpg generator or mand Wright-Pattmern Air Force Dom OH. other method to ive rapid inflation, usually in under Mum 197.
3 sm. Inflation may be activated by the pilot or by stibmetuible valves that automatically trigger the generating unit when it is imumerd in w'ar. A lhird float, permanently inflated, is mounted in the tail cone to provide fore-and-aft stability. 12U4. MODEL TEsMs To substantiate thu analytical data for any typ" of float system~i, a hydrWdynamicaly complete scale modd is tosted in a model basin capable of generating
dif~~eret~~ferir~ '-
__mith
rotors and ballat weihts to permit varying helicopter weight, CG, and moments of inertia. Stability. roll "response, and landing hImpct load data then are re.orded and compared with analytical data. On models tosted to damt, •€•Uoimit correlation bstwca azial-yitir1 and model te &ta has baen noted. With the nmodei in tht taik, wave oorm = W Ivci-ated and mo" tvajoncies .h a headiv, 0o5llin. and
"*
I12-14
2. 77we T~skAe
E&wi*f Jisai. The Timkum
Rollm Dearing Co., 1963. 3. W. J. Morha4d, Loaft Gea V~bwavL Air Feace Technical Report No. 6590, WrightPatterson Air Force ase, Or!. October 1951. 4. K. P. Cokmau, Thety ofSe•-excaed Medwatd Osilatdou of HI agd Rawr BAmin. NACA Report ARR No. 3029,1 &a#lyMemoial Aaronautical Laboratory, Langley Fied, VA. July 1943. R•epblished as NACA Report 1331.
5UUI"*
t., wh "Cuenal"A vih"AtInne fr II*IuWh.
ters", Journal or the Aeronautical Sciences. 5,223228, (May 1946). 6. C. E. Hatmmaond, "An Application of Floquot Theory to Prediction of Mechanics! Instability". Paper No. 13. Rosorcraf Dymmkci, NASA SP352. Proceedinsp of AHS/NASA Aome Specilists Meetinm, Moffett Fid%CA, February 13-1.-, 1974.
CHAFFR 13
CREW STATIONS AND CARGO PROVISIONS OLSmdst US OFSYM 13.
* 13 LIS OFSYMD LS K
lb/ft spiring rate of first qving. lb/ft spring rate of steond spring, lb/ft - lenthof suspessinft - mansof helicpter, slug - ma~ssof external kied, slug l loaed, ftlb/e - aweih ofeterng 9*' fW~5iI eteral oad lbpracice - distance between suspension attach pons ist~c n cse o for-pint ongtudnal X x -logituina disanc in ase f fur-pint
* ~ L Mj
*
rate of suaptasion, lb/ft ~ l~ ~cations to','~
-spring
ML
W *X
su5petision0ft
9
-
7
j
-
Y
13-1
*
U8M i- u ftL
t3
-
-
-
uab awI"&VIA msa5en
speis0f for-pont uspasion ftpit f sspesio sytem Hz - naura frquecy
INRDU TO
This chapter addres the requirements foi pe"sas aocomodatiou and cargo povisions to t incorporated in the detail deign of he~icepters. The disussion of the cockpit includes the pertinent requirements for cmw station geometry, pasmmner compartment arrngemrents. swas and rutraint sY3tans, control/dispiny arrangements, map and data cames cu. Also included is a discussion of the inter face criteria pertinent to the histaIlation and tmployment of cargo bUndfing and survival equipment. Design requirements for environmental control syr.tans and lighting guidelines are Included. The basic design of personnel and cargo accommodations is determined during preliminary design. The detail designer shculd adopt an integrated systems approach to optimizing provisions for personnel and cargo in such a way as to obtain maximum mission affectfiveness. To accomplish this jpoal, the detai designer wil consider pummceters stich as the exact location of displays, controls, in'tsion and emergency equipmtwit turd mutt opti. mize on fte basis of anthropometric date and a human factors engineering analysi. of crew Waks. lit
)equlpewit, -
-
recognize existing demig boundaries w spediisd by the preliminary dr *gn and the appropriate uiySpcfainsPetetMltA)pcfare listed within the chapter where opplicabic. 13-2
PERSONNEL ACCOMMODATIONS
COCKPIT There isa growing body of literature methads, and which provide guidance in the ag'plication of system engineering to the development of cockpits. Howevet, there are few pert isnent, sptcific, absolute roquirenients. The miss ie profile, the vchicle function anplysis, time-line analysis, link 13-2.1
-constraints,
analysis, and tile development of new mauwras anu
components must be considered in determining cockdesign. In additione to design handbooks and manuals e.g., Refs. I and 2 -. reference should be made to MIL-STD-250, MIL-STD-850, MIL-STD-1333, MIL-STD-1472, Ref. 3,AFSC DH 2-2, and pars. 133 aied 4.5 of AMCP 706-201. The general cockpit dimensions, seating arrangements and externial vision envelope will have been deternnined durngM preliminary design. The trade-offs available to the detail designer include standardiza. tion of dimensions and controls versus opportunities for improvement resulting from a particular aerodynamic shape. The goial of maximum vision in all dircsdoni; may compete with instrument panel loý;tion, airframe weight. comfoit, armor protection, and aerodynamic, shape. The detail designer must w.nsider the entire anthropometric rakige of uber populattion for which the crew station is ewigned. The reach and body clearance envelopes must consider the heaviest cluthing, survival, and protective equipment likely tc b worn or used. Guidance for the application of anei~repomeatric factors is contained in Chapter 13, ANICP 7W6201. C~ew sulion arrangement and geometry sAil we in apc-ýrdaitcc with MIL-STD-2S0 and MIL-STD-12333 uniaz otherwise specified by the procuring activity. Deviation from these specifications, or the location 13-1
'-7
*
e
.[V-
of controls or equipmwent, not covered by tiest sped;ficultons, Aus# be based on a huinan factors englossins analyýia of crew tanks. Controls. switches, and levers that require frequent actuation in flight should be conveniently aceusable to the pilot's left bund to minimise removal of his right band from the ~cyclic pitch s"M~ during normal fligt.
trols should be shaped and located oo thai an arua member reasonably familiar with their arraapeeent isable to operate them without visua nrefrec - the so-called "blind position reaction". All controls of lik fuzacton should be grouipmd togtether, and normal operaths w andvergmncy controls should have priority of pasitlusa. Conformity to established custom or standard uunique mi&must be carefully reiwdin cmwere 13-2.1.1 Genera Viale Reqadramests HIL-ST0)4S defines the requirements for air~ '' 'P 'dsg wliin ~ prieont an opportunity for a more nearly optimum crew external vision. Figs. 13-1 and 13-2 illustrate the arrangement through deviation from stanAitofrs equal area projection of the sphere scowigpit sinle-ilo/tadempilt nd ideby-idehelcoper dauds. Standards, in general. represet minimzum wn si gl pl ot/a nd mpl t a d sd*b - eh lc p e q uirements and the result of axp wniewe but a me not inteiided to obstruct pireuess nor to stific initiativo. 'Th problems associated with designing for op How2¶eva the. possibility of improvement through tirnum vision include teflections, glare, distortion, deviation from standard should be subjected to light transmission, and angle of vision. Asitirellective thorough analysis before approval is requested, and coatings shoula be applied as necessary to minimiz= the designer should make full use of human factors neflctions and glare from instrument fune, windenginerin tachniques in or'der to dluternine the seeand windows. Analyses of crew stations under value of any deviation from standa" practie. all part of expected mission proffiki should be conHuman facors enginwintl~ principles also shaM be ducted to determine that transparent areas are adoquite for all mission requirements. Light transanae to deterufine the laton Of any cot8 no MDission
JUAIUW5
UawDpa~Irmla
ma
EM~d o#
ia
mnizud. consistent with mission raquirem-trts. Tintlcd windows or other deign featurcs whier riduo. light transm;..on are not acceptable because of their adverse cf. .vt upon ciiteenal night vision. Maximamn Gbstruction dot to trnpaec fr;m menibars is specified in MIL-STD-85O. In helicopters *ere there are broad expanses of traitsparent areas that are used by more thean onie crew trAwbaur vertical obstructions should be no more than 2.5 'n. wide. projected on a plane normal to the line of sight. This will permit binocular vision to, in effect, see around the obistruction. Such obstructions should be ;ocated to avoid critical vixa on arewand~ to provide maximum distance betAin the crew mnembern's eye and the nearest obrarus:'ons. Structural design and material selection guiomvne f-w tranparnt is iscsse rea inChaper I. hen mirror (see MIL-M-S755) reAr-vision necessary, a should be positioned inside the windshield to Sive maximum visibility to the rear without obstructing forwad viion.Separate 13-2. 1. Coobvise Requimements for size, shape, location, range and dirertion of motion, and foriie-veruus-disphMamaet characteristics for helicopter controls are defined in MIL-STD-230, MIL-STDf"ollwing documents: the AFSC !)H 1-3. Tb. pilot-operated lighting &nd emergency con13-2
*
-
Control pandls, control knobs, handisa, or kG other ihan emerrency controls and integrally illumu.Q nated cointrols should be blackt to wminimze reflections. Control panels lAall be desiped inac cordance with MIL-C-81774. All emergency switches, buttons, handles, knobs, and levwers whioeL require immediaAe corrective action by tht operator in the event of an emapacy should be identified,with altmrate orange-yellow and bMack strip.-a in we cordance with M.L-M-1a012. The edges of adjacent circuar nondoeawt knoias WMl M&%M Y -NLW SL are determined by torque and esoing requirements. Knobs of l[as than 0.5-in. diamketer should be wed for low torque application only. Knobs with diameter of more than 0.5 in. may s&ncvhat reduce chances of error, but iacrease the demand for control panel are. 13-2,1.2.1
Pfteb Control collective and cyclic pitch controls should be provided for pilot and copilot. The characteristics of torce vs displacement for these conttols shel be in accordance with par. 6-3.6, AMCP 70&201l. An a..justable friction device or irreversible mechanism should be incorporated in the pilot's collectiv pitch control. stick sall be in accordance with MS 910M and MIL0-5807. Among the function to be controlled from
. ý
AMCP M0202 thslocation arm lateral aud longstudiWh trm, microphone (radio or ICS)I, weapon fhirig cargo or rescue winh, and caro hook r lease. Subsystem function controls instlldW on the collective control head should be located in accordanc with Ref. 4. [maamph. of cycli anW colle~ctive control grips incorpireting suboystem conftole arm shown in Figs. 13-3 and 13-4.
pedals should be clockwise The force displacemcent characteristics of the pedals .,h~aI be in accordancc with par. 6-3.6, ANICP 706-201. 13-2.13 Seats, Belts, and Haineeues 13-2.1.3.1 Crew Seats All crew member seati;Mail1 be positioned for cane of access to the scat, the helicopter exit, items such as chart boards, and necessary equipment controls. The pilot and copilot seats hAll be easily adjustable both horizontally and vertically. The body contact areas of the seat shall be of open mesh construction or fabricatod from pcrmeable materiak. For helicopters of the utility class and lar~tar, the arrangement of the seats shall be such as to facilitate inflight removal Of an itijuied crevmcnabcr. Seats for pilots, navigators, engineers, radio and radar operators, and electronic
M3.11=.2 Dfrveieag Coatio Pedals The leco tiova and ranges of motion of directions.a control poxain are specified in NIIL-STD-1333. Pedal .sha be 6 in. minimum width. If mechanical ajustmiont of the p.4* it provieed. it should be locite uift- the instrument pawel and rolong the awtarline of the cme station. If a rotary pedal adjustmoint Iia med, the motion of c~mbol to extewd the
AI 0fFSF.QUAL A0t A E'RQJ[C'1%Of THE[ ýPHERE RADIFUS OFPROjOC I E SPHERE £QOALS LIE DECIMETER
Flgure 13-1. Shs~oe Pilix/Taomens-plot Hielicopter Vision Plot
ATF EQUAL ARLAPROJECTION OF Si.t SPwiRt RADIUS Of'PRAOJCTEtF SPHERE EQUALS ONE .afCAIETEFF
ISO
1
110
-
2
--
3VOIO to0-
ISO
ISO
Figare 13-2. SiE-by-s"d Helioper vision, Ptot 13-3
AM~P7GO down strap, two shoulder strapis, and a single point of attachment-releaa with a swngle-action-relese huckle in accordance with MIL-S-58095. Reetrint dovime must provide sufficient freedom of mnovement to permit crewman to manipulate aircraft controls. The releaso mechanism should piovost accidental or unintentional releast, but should also faciltate ecorgency release of is~ured crewmen, i.e., it should be possible to relcase the harness with one finger while tension equal to the occupiant's weight hs supported h.- the harness. It MAl be possble for the meat occupant to make strap adjustments easily, with eithmr4 hand. Ref. 5 describes the development and test of a restraint system to meot the design criteria given by Ref. 3. For those crow members whose duties require tham to suand in open doors or windows during flight, a mo taiming harness should be Installsid.
countermeasure crew mentheis should conform to MIL-S-38093. The crow should be located to minimize danger from rotors. propellers, and turbine blades. Critical dimensions of the seat shall conform to MIL-STD-1333 and MS 33575. aequiremants for struc~ural strength a.nd controlled deformation are given in MIL-S-MM05 and N4IL-STD-1290. Crashforce attv uation shal be accomplished by. plastic doformation of the seat structure by loaiJ-limituig dovioes, or by a -zombination of the two methods. Stat armor may be required. Design considerations for seat aimor are Oiven in par. 14-3.3. 13-2.1.3.2 IW~f aw HamesmUg Safety belts and harnesses should be installed for akcrcw. troops, aud posw qers. The crew restraint harness shall include a lap bNut, Wedstrap, lap belt tie-
SWITCH, TOGGLE, 4-POSITION ON, CENTER OFF PUNCTIO~i TRIM EXAMftEs CYCLIC TRIM CONTROL, PITCH &ROLL
~TCHSWITCH,
TYPEO.'TIONAL SWICHFUSBUTONFUINCTIONs OPTIONAL FUNCTION, DISkNGAGE. APCS* EXAMPUS, (I) WEAPON SYSTEM EXAMPLES, JV)STARIUZATION SYSTEM M OTHER APPROPRLATE FUNMCTIONS, (2) AUTO PILOT SEEMAIL-STO-250 401)SASO
SWITCH, hUSIISUT!ON, MOAAENTMwY FUNCTON, CARGO I'Oc0K RELEASE EXAMPLE, NORMAL RELEASE OF CARGO
I
SWITCH. SE14SI71VIE. MrMENTARY
FUNCTION, RADIO.ICS' EXAMPLES: () RADIO IRANSMISSIONN 42) ICS TRANSMISSION/
SWITCH, TRIGGER, GUAROEC', 2-POSITION ON, FORWARD) OFF SWITCH, PUSH11TTON, MOMENTARY FUNCTION, MOMENTAPY DISENGAGE AFCS STR5.IMAISSE EXAMPLE&SV FORC
4()AUTO
jFUNCTION-
j
WEAPON FIRING
IIS POSITION - IOW RATE EXAMPLE, 2ad POSITION - HIGHRATI
PILOT SYSTEM
STANDARD ARRANGEMENT AND FUNCTIONS OF SWITCHES
*AFCS - Au15:ic Flight CmW~olSPWG$SAS - S'oIII~y Auwm.GRI~f~f SY0110
F~gur 13-3. Exam*pl of Cvewiic Cousrol Grip 13-4
Q.
'
-V
"
capabl of being stowed ianaminimum volum, so as M~ to isseff-a with the traosport of cor.. Troop seats cre to be equipped with bp belts and momddualsif simml be given also to providing shulk hmasne for the occupants of fig forward or sipward facng momts The lap belts should retract OU FNMG COW RTP0ffMONamalil to improve the sm of Ingres and Woes Vhe helM dhoud #A4lt autornatklc* to twrwmiut eat t midm influence. saa The moqk so sine without neod for additional fltklup and hav bees H imar~ly.to J~n equram NuL locking and unalockiing of the lap bel should be pa.b~~m 1 fro point to point wi thbal ow. ble uflt only Oise hanid. ud&atl. vive. to ehsf OCoR*"e or 1 0 IsmLa. er.' ul du useximum cohteht Wm = in the PODmWrbbbb es WA a lower SasYI hl adhwo color scheam " be in accordance with h .1 . me. a- ofaht rru 'IS 74&0.932. hN6e Wmjw~dd muhaiqu whc dhda beeps or qndhud uhei apin a..dllo. to Ptufoe t* Uphdolq and Caipeula 1342 ~ quippe mot ha a emete oal. Vine s. Isf The choice of upbolasrlng aOd carpeting is, to -I-dsempumMugmd W=oP We to be tmuEPO04 WNOMW t"be aiepwa to AMnin90M momsetant. inlon-depondmnt. Consideration shiould be givum to availability. weight, ressance to Par PW .d" wear, sunliht, oils and groom, and wni Sob001buWAoe Dau nrawam adw dwegding influmnoes an drese. coated and hadhe fabric are suitable materials. Foam gu OW =Mmutated aOW e 1h0 mmd. dNA OrLl~uI #Oro rubbe per mIL-R-5001 and polyuretane toam per ' a /qpua. of mas wO ubly, eomst, ued e MIL-F-26-5!4 art t=WccL for US Ary=i-Ym&nsim mvis am~ W1St1a W maYNedtri 14IA " kfmpa Del Cam. Hap sad daue co a be Wooatd in accas*.m wi& MaL4T.5M= so that the men mebe e* reach end Woost tbe maw writho havin to divert Wes a omak. from his Mobs$ "ad mmb.
--
..
dwhe
l-aa
minime mo d comEgardoan are news.-
Ia. hwant or my so detarmle set edinatias L, rerward bliug oul Eme of isgi.. and qp.--- nduu &Auk~eneba is o~inmds by simple, rubblge re. stals qrsle. dad & bskowap of the restrint 4 o ~ maaimt The design must =acdes the reuIOW qfor ame of hasindg and i* W@W" he nouw"a systm. Smuntsd oagma for troop Meatu arn Pr Me~d by Rd. 6. Deuign requireamens for arub-
worthima.and for rearubt )-te oduig. as gweby MIL4TD..WO. Tra"oo em sould fold and he
pficitioins. Seat Cushions shoul
minimize OCCUap
t fttb.
ntarining and dynamic overshoot. Net-type meat cushions may be used if they prevent contact betwain the occupant and the smat pan under design
vertca loads, and if their rebound characterkicst limit occupant return movemient from the point of maxmui daformation to 1.5 in. or les. Cushions should be so contourd so as to avoid constrictions or localizd prussutes that reduce comfovt or inhibit body circulatory functions. Body cotacte Portions of the moat back and seat cushion should be pcrmeabit.
D2.2A Swebq ?"lev O"Irmoabl 1"UIWA9S
~
SCI
Ty W
Cools"L ""WC"
In an area in which amoking is permaitted, me ash trays ame necusary. The mMm-0containedOTC as& trays msut be ftush-mounted flg c m au and profuabl sac..dbl to the crewman's left band. C~mpmtmmas who smalif is sol prmitted Aabe! ~be dedpelsd by an appropefini placad. Crow modoim
411. Cveai Pe. tO0.2a
-lni r governe by MIL-STDmed moin Womb awe disaimd in
Wha u-9 ~d,the bdlatm of the aluminum
Its.. 13-4. Tbrle s( CebDn"
Cmel Hand
Fob %"oftms shul Ne provide. A mmiuamu 13-5
vertical separstiom t'saween litters of Is in. is required. The entire nutalatlon sbould conform to the
streknghrquiri ae-ts of MILS-S496. MIL-STD-
1290 provides design guidanc for providing &deqdat crashworthiness in the litter instllation. Litters may be installed either longitudinally or laterally. Cane must be taken to provide adequate crashworthiness And unless the litter configuration iscap&bis of providing resttaint, the litter straps must be capable of restraining the patient egeinst the curvivable crash loads. A lateral orienrtation or the litters isindicated as preferable. Ease of litter loading and in-flight accessibility by medical attendants arc umportant conjsiderations. A litter lift device isgenerally preferable to adjusable litter support strap assemblies if combat evacuation of litter patients is me
quired. Fixed litter support strap assemrblics should
he used only wher. equired by the procuring activity. The maximum dimensional requiremants (0: aluminum pole Uitters are shown ;n Fig. 13-5. All other~ po elittrs used by the servwce havo the saMe length of pole, distance between polos, and length of
canvas.
-
Oxygen lines 04 storaes for Porn"l oxygen uanits for the ure of all litte patients, medical personnel,
sand other personnel - should be provided. Electrical otttleta should be prVided in order to operate delecical medical equipmenat and provide
tockets for traveling lewds.
As specified by Ohe prowir* activity. foldaway desk, meidical chest and equipment stores. space interons equipmet for two-way commuslcaion between pilo and maedic end a loudspeaker system for giving eamegnc
should be provided.
SURVIVAL MQUIPMENT IN 13-M.. MWgb Esaew en Sevylval Eqnipiene Provisons for iMight esape and survival depend on the performancef capability 91 the heliopter, the nature of the mission, configuration of the Maicop~ter, crew training, and equipment provided. For most helicopters, the autorotative charocteristici of the helicopter end the low altitude at which moat mws siont sic flown make inflight eC&pe undesirabe if
not impossible. Therefore psawbcute provisions are
if$f
APPROXIMATE W'EIGHT, 16 lb LITTER, FOLDING, RIGID POLE, ALUMINUM POLE, FEDERAL SUPPLY CATALOG NO. 653U-L183--7905 Rpgre 13-5. RIgWd Pole Litter 13-6
honstrctione to eveacues abo
*
9
.....
7o062
not required parcifd by the procuring activipnkm ty. If feasible, the role of inflight emergency escape systems is to deliver the crewmen and pass•n•trs with necussary survival equipment - to the earth's surface in such physical condition that they can perform the actions required to survive, evade capture. sa take f.ttion ne-essary (such as the establishment of signals) to aid in rescue operations. Although inflight oseape systems are not now operational, research and experimentation have been undertaken toward the development of such systems. An example of such an effon. is tht; v!EPS program of the US Navy. In this program a pyrotechnic device permits separation of the canopy and rotor system from the helicopter and extraction of the crew if autorotation is impossible. 13-2.3.2 Groed Escap and Dltddeg Plvidlem Beaue of the variety of emergency situations that can occur, the design of emergency egress facilities and ditching capabilities is a difficult problem. Among the primary design parameters arc the nmber, sz, strength, and location of emergency doors; the use of slide ropes, slide poles, and ground
~~~~;
.I..lhinoG~o
...
2
nr
vnnlniwv
mated c exits; and performance features that maximin chances for sueossful ditching. The scope of appropriate round and ditching provisions will have been defined during preliminary deagn (we par. 13-3.2.1, AMCP 706-201). Mission requirements largely dictate the size and shape of the helicopter, as well as its basic performance characteriasfs. These in turn help to establish the natuic and likelihood of the possible emergency environments, An analysis of the characuristics of the helicopter and the emergency environments can be used to cruss awu...s. A..... cri agt"C
,.. ;.U*t
....a
-, _r.L.
£US (UUWU
U36t,1890
n5
emergency-egress provisions. Detail design requires that these provisions be continuously reviewed and re-optimized as adjustments or improvements in subsystems are made. Detail considerations include design of doors or escape hatches to minimize posibilities of jamming due to crash deformation of crew or F.senger compartments. Design guidance is provided by Ref. 3. Included are recommendations regarding operation of amergency exit closures and required markinps for emergency exits.
"{
13.33. Emrgeley Light Pmvriese Emergency lighting pr:ovisions should be installd independent of the craft electrical system in the passinger or cargo compartments. Prcfaan should be given to an emergecy lighting system that is selfcontained, oxplosionproof, wasi-proof, and operable
by a combination manual/inertIa type switch. One emergency lighting unit should be provided at or near each emergency exit. Additional criteria and design guidance are provided by Ref. 3. 13-J.4 Life Rafts The type and number of life rafts, if required, will be specified by the procuring activity. Design guidance for life raft installation directions is given in MIL-R-9131 and NUL-L-5567. 13-2.3.5 Stuiva Kits Army aviators usually wear survival vests. which are personal equipment and for which the designer has to insure weight allowance, restraint capability, and absence of control or reach interferences. In addition to this, space provisions for survival kits shall be provided. 134.3.6 FIrst Aid The type of first aid kit will be specified by the procuring activity. Selection of the location of the first aid kit should be based upon the requirements of inflight access as well as easy access in case of emecrge~nCy Caress. 13-3.7 FIre Extimgishlg Systems ad Axe The differentoa in size &nd configuration of helicopters preclude specifying a standard location for portable fire extinguishers. The portable rim extinguishers should be located in such a way that they are readily acoessibia to the crew member at their normal duty stations. Hand fire extinguishers may be mounted either vertically or horizontally. The extinguishers must not be mounted over or behind the heads of crew members when they are positioned in Ik.r n....... l
SPt~e
ao~raan. ....
5
w
*o-aens n.
Lah
r
rav .
-0
litter compartment should be provided with a minimum of two extinguishers at each cnd of the cornpartinent and near entrance doors if it is practicable. Fire extinguishers for use in occupied areas hAll use nontoxic agents. Hand emergency fire axes shall be provided as specified by the procuring agency. 13-.4 ENVIRONMENTAL CONTROL 13-2.4.1 Vendlatke, Heating, sad Coeling Ventilation, heating, and cooling requirenents are defined in par. 13-3.2.3.2, AMCP 706-201. During detail design the specific environmental control unit must be scacted, air duct routing definl, and the type and location of the air discharge ports defined. The location aid controllability of the air outlets shall be adequate to meet the 10-Idg maximum
temperature spread in both th: henting and cooling 13-7
I o
modes. At least one outlet for each crewman sdhll be located to provide variable airflow over the head and chest. This feature significantly increases comfort in the ventilating ard cooling modes. Detai; design guidance on environmental control and air distribution systems can be found in Refs. 19 and 20. An example analysis of heating and ventilation requiremerits is presented in Appendi.- A. There should be a comprehensive investigation of the possiblc toxic elements from all of the materials that go imo the construction of the aircraft. Considcration must be given to the removal and the dctecuon of any toxic elements which may enter into or be generated within the crew compartments or the cockpit. As new materials are developed for use in the construction of the aircraft, they must be tcted to detormine possible toxicity. Since it is impossible toprsmnt a comprehensive, up-to, date list of toxic agents which may find their way into the cockpit, it is noees"sarythat each design program consider toxic hazards. 13-2.4.2
Winsddded Defogging and
delddg Equippm The windshield defogging and deicing equipment
,dmuid 09; pJUV3UIGUQ
t th© "quiMnt
o
altitude, thermal. and weather conditions required by the mission profile of the helicopter. Detail requirements are given by MIL-T-542. Automatic and manual override controls should be provided as rppropriate. In the case of anti-icing through chemicals such as alcohol or other toxic substances, the design should be such that the fumes or liquid will not enter the cockpit or passeng compartments and thus advrsely affect the crew or passengs. 13-2.4.3 Actnmdeul wFrem scisl The noise within occupied compartments shall no! be in exom of the maximum allowable leveis prescribed in MIL-STD-1474. MIL-ST.?740, or MIL-A-8806, as applicable. MIL-S-6144 provides the general specifications for the soundproot'mg of aircraft. Speciil attention must be paid to tehlmiques for reducing noise level at its source, and to the use of special materials and techniques for insulating against acousticaln ise. leference should be made to par. 13-3.2.2, AMCP 706-201, for information on maintaining noise levels within accptable levels. Auxiliary systems that normally operate for longer than 5 min should not produce an increase in noise levels in occupied compartments above that specified in MILA-8806. Special missions which may require noise leels lower than those required by the genetal Military 13-8
Specifications should have the requirements so stated in the detailed procurement specification and special care should be taken to insume that thee levels are not exceeded. 13-2.5 SIGHTS AND SIGHTING STATIONS Provisions must be made for safe, efficient positioning of sights and sighting stations required for direction of helicopter weapon systems. Design guidance for both direct and indirect sights follows. 13-2..1
Drset-ewi
Rgs
Direct-viewing sight may be fied or slewable. Fixed sights generally are provided to permit pilot operation of fixed weapons, flexible weapons in the stow position, and rockets. Fixed sights should not restrict pilot movement or field of view during normal flight conditions. Furthermore, the fixed siht must not interfere with emergency exit crew movements. The sight may be folded out of the operator's field of view until required. Sufficient clearanoe must be provided to permit the pilot (with gloves) to place the tight into its operating position and to perform any requiired hand-operad r"cCd1 adjum 'en•ts. A direct-viewing flexible pantograph sight may also be provided at the gunner's station. It generally is supported from the aircraft structure and is hand directed by the gunner. Alternately, the sight may be fixed to, and rotate with, the gunners seat durbig target tracking and firing operations. The helicopter designer must insure thut adequate clearance is provided for the gunner's head, hands, and other extremities to permit unrestricted movement of the .ghting station installation. Airframe and cockpit surfaces within the operator's f. iew should have dull (antiglart) finishes to prevent reflection and eye strain. The canopy enclosure(windshiold)withintheline-of-sioht envelope should contain as few areas of curvature and thicaneu variations as possible to minimiiz opticad refraction and distortion. That portion of the windshield within the sight-aiming envelope should have adequate duefogging and deicing provisions. Direct viewing sights have been a major source of serious or fatal head traumas in crashes in US military aircraft during World War 11, Korea, and Vietnam. It is essential that all sighting devices be designed to minimize the potential for injury. Safety factors to be considered arr 1. Capabilit. for instant removal, jettisoning, or storage during e eragency periods should be provided. 2. Stowed tight positions should not create ad-
K
• dlk~ai besart to odll• aew persod in dow smt "~3 ',
"
of "mmen
A deso
.
f stwe tisdon sUvrmth abould be
L%.
Mbmlk rlmba Sbmlm sai
for miWW in Imtaa ltoapow-
wWl eMploy tM• do~t iustalkaUs usil forogum ad
pRovidd to prevent the @Wt fmmr bodidring
rockes. Provision may be nqtdrW for installation
13043M
obtained slectrically or mechanically, and enerally requires provisions for adjustment.
•L 4. Jettisoned or Nnoved slghtas should not rupremt Ithal miasle hazard in the event of a crush. S. Thi guidanew ontaiud in Ref. 3 (Chapter 6) should ha followed.
Hdm
Memu S
The helmet mounted eight is another type of direct-
of reratabl op6W fltar (in tiw gunoa lin of s•ht) for e whe the misi exhast liot intnsity is ufficiet to cause temporar blindness or a blurred view of either the targpt or tde guldpow refWmNu. Sight alignment with the aircraft datum plae may be
v i optia diht curmesy used for heiopter appilestios. The sume cerane coiderations app-
13-3 LIGHING SYSTEMS
ly to this installation as to the aircrft-nounzted Naebi SlglOL Adequate hend vlearance must be provided betwee hem et pn eCth &WAd the Glctft structure for n•nora flight opeation and for wGIon control within the mquired azimuth an evation flexibility ramuu. Quia-dieconect provision must ba mude to slow dislodging of the helmet from its electrical or
13-3.1 L KI• Ol LIGHTING SYSTEM The exterior lightingisystem design may include an anticollision light system, formation fights, landing/ taxi lights. searchlights, floodlights, and position lights. These are discussed subsequently. MIL-L-6503 provides specific design requirements fw exterior uigbting.
Onuchealaml onectiou with the aircaft; provisions also mum be made emblinS such disconnection to be oanraina. P' eight IooMt-" SSaOne typeof o elie incrprtes a medhani-, OW linkage to detect ead motions and sum the aiming informaIn. Other ts of helmet sights use ""glat senora, sonic semor and elec"magnt fld to perform the asam function. Anotber type of diretsight is the periscopic sight which views the tarlu thrcugh a mirror arrangenemt and permits the eight bed to be located extarnally and in locmtio that allow unobstructed linr of sight within a large envdope of azimuth and doprIeion ng limits. ThMoptical equipment may be spported from the aircraft structuw or may constita to integral pert of rotating gunneW's station. Peico installations of either type anr sizable and require adequate airframe support structure. Power requirements for a flexible gunner's sight station may be significant and may require early consideration of sighting station cabl or hydraulic line sizing and toltli,
/
1
3 ludireet S• Indirct sights receive te targa IaWge from electronic sensors and project it to the observer by means or a pael-mounted or helmet-mounted display. The ability of the obeerver to detect and identify targets with an indirect sight is dependent on display resolu. tion, contrast, number of shods of rey, display sIze, and eyto-pandistance. Detail design guidance for -'Indirvet view sights and other CRT displays is available in Ref. 21.
13,3.1.1 AndediUea Lijgm Sys•m Unless otherwie spec• day/niSht anticeollion lightW., syste sArmnybe evdaopei provided. TUi system is idetified as Light Set. Navigation (AADSHIL) and provides a white daytime strobe of 3500 effactve candle-power and a red nigbt time strobe of 150 effective candle-power. This syu cm is available in the following two types: I. Type 1 (28 V) NSN 6620-00-361064W 2. Type 11 (115 V) NSN 6620.00-361.0614 Field of coverage Aall be as specified in MIL-L6503. 13-3.1. Fermanleai For helicopters, formation flying lights usually fall into two catqtorioe fuASdg formation flyig lihta and rotor tip lights. Fuselage formation flying lights shahl be so arranged that adjacent aircraft can fly in either steppel-up trail formation or vat formation by alignment of lights. Durable elctroluminescent panels easily can be adapted for fuslage lighting. Refer to the specific aircraft system specification for detail rm quirements. A test installation is required to verify optimum system design. Rotor tip formation flying lights should be con. sidered on helicopters because the rotor blade usually extends beyond the fuselage, and movement of the rotor disc provides an indication of an impewding maneuver. Due to the high centrifugal loads of rotor tp%. multiple white lampe should be used minimxizng filament and vibratory angle. Some type of 13-9
slip ring normally will be required. as well as wiring in the rotor blade itself. Five different light intensity levels have been found satisfactory. Formation lights must not be visible from the ground, and should be further shielded when practical to be visible only from behind and at the naow altitude and slightly above the lead aircraf. 13-3.1.3 1.amihg/Taxl Ligt A 600-W or 1000-W retractable landing light (lMO-W preferred) shall be installed. For small helicopters with limited power systems, lights of lesw wattage may be acceptable. MIL-L-W53 requires that this light be slewable from 20 deg above to 60 deg below the normal level flight position of the si:craft. For those helicopters having limited lower nose space available, the 20-dog light above normal 1%;vcl Rlight position may be, difficult to meet, and a deviation should be requested. 13-3.1.4 Searcl~h A 450-W controllable searchlight, in accotoance: with MIL-L-6503. shall be provided unless the dcta~i specification requires a larger searchlight. i.e., 600-W or 1000-W. 13-3.1.5 Floodlghi System Some detail specifications for rescue helicopters may require a floodlight system in addition to the landing lights and searchlights. The ground area to be illuminated and the helicopter altitude when using the floodlight system should be determined prior to the design or during a lighting mock-up. 13-3.1.6 Position Lights All helicopters shall be equipped with fuselage side position and- tail lights as dethed in MIL-L-6503. 13-. INTERIOR LIHTN The interior lighting system design may include cabin and compartment lighting, cockpit lighting, panel lighting, interior emergency lighting. portable inspection lights, troop jump signal lights. worktable light, warning, caution, and advisory ligh's, and instrument panel lighting. The applicable Military Specifications for the interior lighting system are MIL-L-6503. MIL-P-7788, MIL-L-5667, MIL-L27160, and MIL-L-2547. 1"-.2.1 Cabin &adCompartment Lightlng Cabins and compartments shall be provided with suitable i.Vht~ng for passengers and crew. These lights shail be ins.'Ued so that their direct rays arc shielded 13-10
from the pilot's eyms and so no objectionable reflections are visible to thes pilots. In aircraft where dark adaption is required, these lights All be caps ble of providing both red and white Wlnummatiua with separate dim controls in tWe cabin arma The required levels of Diuminiation are tabulated in MIL-L-65. 33. okitan Cockpit lighting shall provide illumination sulficient to en~able crew members to aswertain readily indizators and switch positions. A cockpit dome fight, with controls acocssible Lo both pilot and copilot, will normally meet thi%requirement. The dome light shall be dimmable and provide either red or white lighting. 13-4.2.2.1 Utility Lloa MIL-L-6503 gives applicable dasign requirements fo, cockpit utility lights. For most helicopters, one light is installed for each pilot. 13.3.L2.2 Swmedau Lighti Socont~ay lights shel be instalies in the instrument glaze shied to provide disumable red and white iilumination for supplementary and thunderstorm ligh1tins. Th=s lights, coiuec-ted to the mactis!W bus, hllbe in accordance with 141L-L-18276. Utility lights may suffice as a secondary lihot sourct in cottamn cockpits if they can be located to illuminate ca sential inatrumcnta while remaining readily accessible to be used as utility lights. 13-31.3 Panel tlighti Control panelks shall be sufficiently lighted to permit easy and accurite reding of the information contained thereon. Integrally illuminated panels shall be provided in accordance with NIIL-P-77g8 when a red lighted cockpit is specified, or in accodance with MIL-P-83335 when Air Force blue-white light is 13-3.2.4 Interier Emergesicy Light An interior emergency ligting system when required shall be in accordance with MIL-L-650. Design innovation radioactive luminous lightinig panels may fulfill some of the requiroments of MIL-L6503 for emergency lighting. 13-3.2.5 Portable Inapectloa UAWht MIL-L-6503 requires that each helicopter be equipped with a hand-held scanning light. The light covered by MJL-L-7569 is approved for this application. Outlets ihall be provided in the cockpit and crew compartment to pcrm.4 the required inspections during hours of darkness.
-
)
AMCP 706-202
113-3.2.. Troop Jumap Sigas Uph Midefined! in MIL-L-6S03, when requirral by the procuring activity, r troop jkui'p s6*01 light "I be l?Tovided. 13-3.2.7 Waruh, Cawdee, aud Advlwoy Lights The warning, caution, and advisory lighting system is iiiscussad In raw. 10-2.2. 13-3-2.8 laubuuwst faiel Lightg Instrument pwnel lightingl is discussed in por. ) 0.2. 13-3.2.9 Cargo Computsuei Lkgbtiig Cargo compartmnett lightang should consist of, at thn minimum, two rows of flush-mo'ntitd ceiling floodlights located along th.-ruges of the ctiling. It also is advisable to provide "xternia ligaitir4; in the general area of the rarer and other doorways, to facilitate night loading and enhance safety. A row of lights along the base of the side walls provides the il.. l'imination required for the r&iong of tiediowns on vehicle frames and undercarriage, pthe
13-4 CARO PRVISONSThe
-~~
*
Any atructurcs or components which project unavoidably into the, comnpartmcnt must be marked ;onspicuously. In all ceaes, the protruding comnponent must be suitably protected agailhrst impact from cargo and vehicles. With the ovc;rcrll dim.-nsions of the cargo floor determined by prelimin~ary design, the cargo tiedown points can be located. These points must be arranged in a basic 20-in, grid pattern. Such a gridl pattern has been standardized internationally (AFSC- DH 2-1) and is shown in Fig. 13-6. The requirement tha cargo tiedown fittings be located on 20-in, centers must be considered early in the structural design process. since this is a major factor in the location of fuselage frames (sec. Chapter I) Helicopters which have the capability of haul:ng vehiciet should havc rtrengthened treadway areas on thc floor, locatad to coincide with the wheel locations of all A-my vehicles which might c transported by hsciivopiar. Th6 auj isUQ~ J1Disciassc4"AIin CAMPM a`- caro dor
houl
hae enap11.r
equal to the sixe of the cargo compartmenet, with the do IIo h aepae stefor ow.do The provisions of this paragraph are applicable to heliopt:~ avin anallcaro orcor bied ft doors should opets clear of the extended wall and argco/pass. hagewaringan ill-aion wihrh cargine ceh~ng planes, to permit straight-in loading of the highest load which will fi! into the com~partmnwrt. In carried within a fuselage comnpartmnent or within a basi heicoper irfrme. addjlion to the main door, at least one smaI'kcr, -vconweuý;ite rdo eloseoa Mhrd dar doo provisions, cargo oprnt.incidental Bagg the cargo compartment, ýoallow acces when the and equipment stowage bins shou~d conf'rrm to te coi~pa~ti~n cricri w1~vcrcomiparm, ifilled witri cargo. If the second&ry WrLsa5U door P- on thca side of the fustlage., it should be sible. located on the right-hand side of the helicopter to Army cargo can be grouped into two classes. Clas~s facilitate the pilot's surveillance of loading A ~~~cargo includes unpacked items and loose boxes oprtns smaller than a 3-ft cube which can be restrained by a If required for a cargo hoist (Chapter 11). a hatLý net or similar device. Class B cargo includes larger, atlstOinsqaehodbeocedntecner singlo-unit loads such as vehicles, artillery, and fuel line of the floor, at the approximate location of the barrels, which are secured individually within the cargo compartment. Incidental cargo, which falls hlcpe G into neither claws includes items such as spare parts. 1,34.11 Detail Desiga flight bags, or mission-oriented equipment carried onboarti by passiengers - e.g., tool kits. ammunition The cargo floor should be. sufficiently flat and boxes, and weapons. smooth to insure that boxes, and fabric containers can be slid across the floor without being snagged or 13-4.1.1 Cargo Coapmartmew Layout damaged. Tiedown fittings, floor panel fasteners, and The basic enavelope dimensions of the cargo cornseat/litter anchor studs must be flush with or recessedl are established early in the preiminary below floor level. Tiodown rings which cannot be in* design process. The detai desiger must keep the stalled flusii with the floor must be made removable holicopter structural wid miechanical comnwhen r,:' ii' uwe. Since they pireset a storage probklem \pouenas froma intruding into this envelops rA the. and r-.. i; Sciaei &Iinoonveineicae, removable tiedown rings should be considwei~d only as a last resort. 'tgldevelor-. Compartment walls %nd ceilings 13.4.1 INERNAL CARGO
A
must renmain unobstiructed; even a minor protrusion can decrease the usable ctrio volume considerabj.
)pertinent
\~vatious
AMW
M20.(ALL DIMENSIONS IN INCHES) -. " 10 20
', I
• -20• -o --20,°
--20 -0 120•'
I 0
S/
.•2
201.' 2D-.4--20
20
J-t
o00o0 o
0
I
I
o00
I
I
I
0
0 II
0
'
106
o o
I
-
,/,/FORWARD
4s"
I
I~
i
6
AFT BULKHEAD
BULKHEAD
EXAMPLE -TYPICAL CARGO FLOOR LAYOUT ""-TP
n'-O--9 Tn 110n
-D-.70 TO 90
SEAT
15-25 r4-j
-o))
.•,.,I
IS (REF)
/t0 D-s-0-oTO D-35 TO
I
REE 0-7WREF)I
.2315(RE
IE\..... I, 0 ---
L C•. . :-t
c
!1
IsI_020
iin)0Jmmj10
2015--2
20 20
SECrION A-A ELMIATD.FOYATED OFF]OO AD R~• 90" 1. IF C
c~* 6]0 FUN
FITTiNG
CFSHALLBE A"AFI•TTND OF SEAT LEG
"LO"ED20 OUTBARD
16-N25 15 E- 0-
OCA-.
BYP0CIt
B
LATERA.L GRID PATTERNJ DEVB..OPIEN1 BETWEEN LONG!TUD!AAL LINE O, SEAT LEG FITrINC(5 500 POUJND FIThiNG LOZATfJD BY PLACOAENT
ITNSSALe 2. IF 154C<20, FROM INTEFSEC VON INBOARID LOCATF0L)1.5 FOOM AND FRM 3. IFC)
T2
CITN{{15
ELIMIATED.DETAIL
4OF
0
20
41-0
5
0
ALLS OF T0PUNSEAT L 5000 POUND FITTING LOCATED BY LAI5IAL POUND FI 7,1rING LOCATED FRM Vl2V A-A I10.000
Flow* 1341 S6Rmkii Caw I.• -12
k"
Ti'sw. GCdi
AMCP 706-=0 Provision must be made for the lnev'1table damage cargo compnipment. Any such components locate' whic will occur occusionally to the Roos during iinmecistaly adjacent to the wall of the cargo, comnloaiding operations. The Amo should be divided into pantinent must be 9Akiled properly to guard against panels that cmn bm replaced readily during orgenithe poaability that shifting cargo could deform the zational mainteniance. If it is tot possible to make the wall locally and jami the controls or damage the comnload-beariol floor ssruwre rnowable, tien it ponents. dhoul be protecWe from minor damage by rtplaceýabecovers. The fAoo should be a&&4 w stiff and 13.4.1. L~eaft AWds p~tc~r.reutan - ossblec~f~l~entwit we~bt Tb. detail design of the cargo cotnparmean mwust limitations. Material meleccion should te based on icuepoiinfrtebnfn fcrodrn consldmatlons Of wuight fifss a mt"@, to k*;A prvswns anlde frThe utndlitg oflu crof dthe corrosion. fire, niso~stwe, and abrasion. Deamust of lodg dswaeTh utiy 'w fte tie einuea~treq~ut m~to ror ~helicopter wrill be affected greatly by the attaition thik aimreceies fom
cost lso s B*fbmn dimshoewdpay Patcua ateai otkM tal e~ecwfcoronon -n Caldbmof orrsin. admlis w~j~ntsin he vicinity ofte tiedowli AISIA should be dtW* Carufity, wspcially what the fayiaS awl %"*samre ub~Tb ~ jogtorocowivs formw The floof wtipcsw ium o isac
deesssgny.
ne
ramp should be at least equil in width to the cargo cibinpurtment floor, and should rirovide a contiflUwqs. sin0(Dth sirfacm over this width, ahthou~h it may be mignem Wd inok left t'nd right halves for better
con'km~ace to rough terrain. It sho-i'd hav'- the capability of being adjusted to and locked at any within the widest possible limits. Measured _ h _t -.. -*d& we -0""%Rm" '6 "'1'--.fren W!NOWatk wfies depim.e odn ievel ground., minimum range of travel of -10 in. to +50 in. is f the aevct. ross an abCsolnccuulypteed , andm "P C~Uk$UtQJ4 he i~tif5 CatM5 cosirecornmnmded. The slope of tie ramp when deployed J =M t F4n an Un fo tforA h on level groianJ should not exceed 13 deg. Allowance mateialmus beh~hy ~sl~t t waer igration sliuul be made; for dxangc in bdicutipw floor height furifuaM attck. The looe shold ~ itha ~ duwini loading. Jacks supporting the helicopter nmb loorshold to~J * a nsisip w~t wll limiatcsuch se~hng. but atthe cost o nialetia. mitas teType MI mu*thI decribeiSi cruide~rable **St. Alternaively, a lo&O limiter (we MIL-W.5044, to protide good footing tv puznaf~d Chapter 11) should be incorporate into the ramp acand o0442 u&ction for vehicles. The corro mamtuating mn'chan~as, to allow the -amp to move uppetamovit wulL and ceZMdi soud be biad with prowad~ as taquijed when the helicopter bettles under tocivepwwing Tis materialsboulda be a light n o. Without such a !oad liwkni intalled, the ramp poiAble &and sobotk sve: to pro~t4 the ft fu@ap actuatoh and it. supporting stinctaxe must be utruc.*re from muinor damage and wear. fr n'1sM.desrgned to support the weight of Cm loaded belioutr and awusid hwnsution, sad provms a IMOGO.l, snm-g-ftee wall. A comnucierally avaslabl FibergaCargo helirpbesn saus have povwer-operated reiiforcied Laminated siim~t sntenAi Mneed wO"el 111 ramps. Power for operating the ramp and associated a cargo comipantmelnt hoer. The thin Usweal is corkdo-"s mstW he of a type availabie frum pround power shr'tec of paralki gias fibers bonde together it. a units. Ligit obsevation and utility helicapter may croweply comnatrion with wmoy samin For adus manuly positionod ramps that can be reinotely f'tional insuilation, FiberlsW battin mateial cao. be notred within di.. carg compartment. itea entrance inWaldW betee the MWe Mmwftmssat lIner and rap muct be deVoyabe in fRiot. Aniy doors the heliecopter skin, operating in conjunction wihb these rm-ps must bW ALI materoal used for cosstrutam saw lhnit the caipsh of being opemsd in Rlight; alternatively, such wmgo coinpaitment mwut be fire-resiistant. Any cargo doors may be ranovable on the ground, with their imnpuhtumt if notc r 101111' c~ nfight mut aboetce having no infact on structural integrity or have~ a fire dsteatcv aWoremnotsly controlled fire exflight caatrsis fiftinuising squipawnt. I. amy car, a mean should The ramps should be strong enough to permit bc1avdeto sea off the cockpit ftovm anoke and loadinig of the besisiest anticpated vohicular load. fue riginatiug ia the carg compartumet. with only tint corna of Owiramp contactinj the Mlight controhi, crkical musehanical eommponnts. ground. The ramp miqy be ttcgh'ndkcally to vrd critical wiring and plumabong mumt not eater the provide wnedway areas at kaw equal ioi widthi and .
)
Ithw.ual
13-13
4
struagth to the treadadays Withf the Cwg compgartmeait. It ks dedrabse, however, that the entire ram~p have a ,unifona beas*n eureja#V1 equal to that of the comipar~ncnt truudways. Ramp edges and dorr sills motI be designed to withastad the uvze lcaizcd impact koaiiags encouatered iii cargo os"ations. Li. the absence of mome ipecfic design ciitcria, the ramp aftes, door sells and vertical door frafms should W3 domigp'd to sustain without damage a single randonaly loo- ad load equal to the weight of the heaviest single item to be loaded, the Woo~ bng applied to the structure through a 1-in, radius sphere. Ramp ejiteusions, if necesary, shoulcl -vnform to the straigilh requirements of the integr-al ramp and compartmerit tr~adway. Extensions should have a Continuous width equal to that of the ramp. If this is not possible, the mort narrow individual ramup actmisiona should be made reveaible, with a wmnooth smiface on one side for cargo loading and side %inonon the opposite side for guida=c of wheeled vehicles. A winch should be provided for th pups of looding aad unloadin caro fivim carow helicopters, Although the winch should be locaed at the forward ew4 o~f thc carga o -rtrnt, it. m-- be locatedL. elswhere provided that a suitable combination of bMocks and pulleys can be arranged to guide the winch cable. As a minimum requirmnt, tht winc should be capable of both pulling onro (,n boa'4 and extracting cargo from the compartment to the ramp. Reveral of direction o" pull caia be accownpished by rrgngthe wioize cable over snatcb biocks located Oil th ramp. Snatch blocks that can be attached to cargo fiedouwr rftuio will gieatly improve the fleniyaut.-it V wdnessbfary M cksandpulley tihe part of the bauc eq, kimeat of must be prc vWdad the helicapter, and provisons for thtir mounting wnd bsorege must be ifieAded. A &Wsrable, but nonauentiAl secim!-ry mode of opouaion for the winch is as a hoist witbin the cargo compartownt, with the Cable rigged ovar a ceilingmourned pulle. If a floor hitch is avaiiable,. the caro winch can be used !o a limited degiee as an eateiaa load hoist. clowee, the designr should consi-jer the nounwung requirummis wbich this avplication impac,:* u'e fthcarm .%%ch. A winch used only for ca' .handling andetar-oave highspeed capabfitiy. Thus, a'tbw4h fomc imis an quite iigh, the low cubc aprqd tends I. mimnmmnz powNv requiranenta, Wim, and weight. OL, tine other hiand, a', Imina3 load bmth. tc '-- of any vaiue. requime a isa reslin rot uai.~ %d, as w"l n a rapid depl~yO'Vi-w rats wnd A*h features as a,bb-capacity bra ins pyrotechnsic cab" Cutters. 13-14
rTe minimum, ueLile able leugib m"s be that required io rstrisim a lond wh ishI 20 ft bmyon dt aft-most paut of the belicepwa (the tPall otr dlk. or, in ;he case of the tandem-rowo helicopter. the aft main rotoc tip). The winch systom should hae ourWhait capability to hvAu a flat-bottomed packag of a weight equal to the nmaximum peyloid up to the loading ramp wnd into the carg comnpartmenmf. on tfr3 auseptiop that tde peackg is to m~ita skids, and the ramps deplYed cia level ground, #he ewe of "~tch blocks and multpl purchur cable arrazteemeata is recommended to achieve this maxiload capability. sinm a witch with a sriaigtmumw pull capacity of such magnitude wcald be probibitiey large. The oof~ueffIcsts Of friction showil in Table 13-1 should be nad when adetrinling winch capacityThe seleution of a power Moturv fo the cargo winch shoul take into comsidaradoio the fact that kondie opratvions uasually are conducte with the helleapta main power plant shotI down, and power is supplied by either the onboard auxillbary power unit at it g~round electrical power cart. Conro of the winch .0.
ueh*~~mZa
.~.
long eIoSgj to permit the Wis.. ovewortot 10
0
th~hu the Cors comPartment 46-1. ramp maybc u32l~seto ecu cariU wtieow dermy Whic myb sdt euecrowti ryhl copters. Floor tisdown fittings must be conpatibl with all thee device 0-ovliioiis must be made for fte storag of an appropriate number ot tbue doVIQU within the Cargo Cooup~entut. 134. UTNAL CARGO Th onco-sove? practive of cA~yirgcarso lo"d ex-7 zYV;~ "ffu~
-A*'wvs
mu
m.v
a st~andarfd ze~rtlg pmceduma~meecally in cowbet oparalAu. In muan owns, the usigo is g~iftTABLE 134. (-F1Ck
. OF F1A1OVN
NATERIAL1
5( ~E MIS(OFORN eOW; ON We~t ONM A METAL ON METAL
rur.
O RC0
h 013 TO 01~0 07 -0.15 TO 010 -
0.08
TRALKED VEHICLES GREASED SURF-'.CES
C.05 TO 0A
WHEE..ED VEHICLES
0.03
[AMCP7022 j
~TABLE M32. STANDARD CARGO
TIEDOWN DEVICES
DEVICSIZ, (ITYPERATING DEVIE SIE, t TýPE lb 'SPECIF ICAi ION
9X9
A-Z
MIL-T-9166 10,000J
15X15
MA-2
10,00
ISAO
MA-S
10.000
CHAIN. TIEDOIN
9
C-2
10.009
WH**-44480
I IEDOWN
9
1MB---
10,000
MIL-T-25959
5,00 5.000
MI(A-) MIL-T-7l19
MET. STEELLCABLE
MET NE
i7.wrING
NET, WIEBBING
4CHAIN,
STRA-PTIEDOWN STAP TiSWN SRPTIEOCIVN
-5
--
l
I~-* S~RAT*D(N 15 A-IA 1 20
GCU**iB
5.00
'A
I-
(A)OMI SINGL-E f JINT
4I-2 7260
(B) LATERAL TWO PCINT
A *~off
t tama.m I)
5thdkcplen
loaded froma helicopter evn though it could easiY fit insid fte cazgo comtpartmwnt of the sane belicopter. Numerous advanflgw to this method: I~~.The helicopter neaG no land, either to pick up to release its cuggo. Loading timeo ix minimiuid for the hookutp and
)2.
pare to hostile fix. 3. Omimnia cargo can be carrie. 4. When wing singepint W4Cem, loading I and unloading have liak or no effect on kmtweadinai CO positicut, and thereore reompatcion of CG lwtoaor is nat stiuin 5. Cargo cmt be jettisned to lightec the bellcoperT lod prior to an mnmpvay landing. G,_mall, wason examMmrp cnr~vu-(1)) ha~fm can be clasfOc by the nrntb's of points, thrueg whicL the load is atbclAd to the ehelcopter. fw-pin, iurk~-point rnapww'ins n(os~k V FueFg. con*W~M h las n0¶v4wity of
(C) TANDEV. TWO-POINT
FOUR-POINT Fixer 134V rxkMruS S11yuvalmmk, Cnflprath.
Bak
Most light belicapmer and attain early cur onm becawe of 6et av model sachien singltt-piat caqibilty by usn a usardfability of cantafrline hard points on existing shofl foeu'msmber stag the lop' of whlk*' ama aib4 beause few hseas. with the notst~c anduosud to am sape brw tde COD ofls9s,66s esoapdwmdmof artilley pwain, are coat ruud foethre The carg in suqded fois Ibis *MR. While this arpoint pkikup. raqpmmet wnw to amed the Imi iota the aLla a&Mficm to the disunion that follws dsign fran, allowing grnupid stuvctUral mitt, it has is guidance for eneraal load syar is pro-v~d by fe of xmor the amupeica undesiabl si6& Rd.7. point too far below tur CG, with tin rmalt ibm in Singlo.1cAnt amg~vloo -e by far the simplest Flig. 13.9,A). nmesa of canryin mgwarns cwpg can be canMe As Ith kiadJ swimip hiaofy md loagitdinaly, the in a net or on a Palle, with a mniahww of psups'A vsaor pivots tkit she siUPeiOa pout at the0eargo ILLwe ratio.. Theft is so mud to pack and Cmma~uly MMg Nig remsrzit facwos a mnthe casttz: &PeK, impoeig Sadamkisl WFUnUti amournS upsm thu hioeptw as it duierp widely (ra the with load cauied within the wage cvauputmuat.
1
d
d P --
M-Pd
M-Pd
SUPPORT STRUCTURE
jik LEG'SLING
MWFOUR
PP
P S(A)LOW
(B)RELOC~ATED SUSPENSION POINT
SUSPENSION POINT
d
d. M-Pd
12-Pd
30IRTAI&Il
/1
I
/
~SUSPENSO
POINT
TROLLEY
--
~VIRTUAL
SUSPENSION UVDPOINT BEAM
LINKAGE FRAMtE p~Pp ~
Ij.
I(C)
(D)CURVED BEAM AND TROLLEY
FOUR BAR LINKAGE
Ulpm 134.
t
d bW4 do Som ofai
Poft
4-..
-
.- ,-p•
--
m
•
u•
iI
lI
AMCP IMM~ locstion of the CO (Re(. ). The mokhios to this unstable ituation is to raise the swpemiom point to a locatin M dos as posble to the CO of the empty hliooser, This can be doam by physia9 marranging th strmu a shows in Fig. 134-(). For mo" cume inchaig main ellco" , it is mr kmM to crese a vrtual supesnnoe point by andhoag the cupeo book or peadent upon a linkage M&fg. 134(C)) or a alUy naunings on a curved track. whiah allom the pedant to mOVe as though it wom pivotn About a point nmee the CO (Fig. 13-(D)). "Themajor difficaltm with simgle-pomt suspmnslow arie from th fed that te s•nso can provie o01Y simple Pdulr stability to the load, and cannot provid any rraUiw or stability in yaw or Pitc. An auzlry lif onunecting the load to the + opke would provide ptc and yaw restraint; ho-ever, this iored lis cam"ot be used becaus it woldcomptie the m Pe mim capability m34 moment upon the poMMiby impoSS uNcontWolla helicopter (Rd. 9). Thus, the oWy mm of providing stability to a single-point lod ias by using the a.rodynamm fovmsemermed by foumd flight A nrachuftainaflh I tohie tgaikjn amu nf
K
) csoW bmca prmvie the nmeucar restoring nu)most to kWe e= otharviss unowabl ielod
-t
where W - weIght of aderuallOto. Ib x - dimance butwenm suspeaac attach pzrnts ft L - ngth of suapemnso•ft It can be sm that s"tear saspemmon 0ablos generate a ete rutoring moment and result in a oe ,tabmload ad highe allowe elapdL Pitch rewtaint is provided by the tandem type of twopoint suspeanion. with the distance between the hebcopter attachment points detrumining wbether the load pitches up, remaidn lWvJ, or pioch. down a aerrdynamic drag swig it aft during flight. To avoid having a low-dnsity load ily" up into the helicopter due to a drag-ioduasd pitchp, it is dcsiabne to have the Ilod attachment points spaced farthd- apart than the helicopter attachment points. Lala-ally disposed two-point smtmumns do nwt provide any pitch restraint, but do orer ome roll rmeraint as well as yaw restraining torque Four-point suspensions povi& simultaneous wrstraint i ;th and roA, &Wprovide a slotly mrU cetive yaw restoring moment N. as expesedy a1-2£dlb
aligned with N Ik P tlii 14 the diawio . of abL Ha• or, the drwp mo-s its N !.(). +7.3 LWt,_+ eafct A very low s ,eedad in hovering flight, wbwe7 rowa d"wokaa can ajapy comeidw"bl rotationa kfous. on outain types a ioaI l - longitudinal disance betwemn cable attach Tbowa li•l•i amountsingcident of yaw restrint, IfPnt ofier a fldwd bwWswma W ya reaw t. i Y -latera dietance oe abavc a hook is rigidly moumntd to the airframe at the sue-t ,, supnebyasotoisf b load ponadif tAe A four-point suspaenson is compatible ith most bya ' o'dIuP multiklegding wtkbaringatikeapx*tber~g inuw" v Q on r Wei MW pnmits ,.,,,...,,, ,,..,,. Wv SWv-,, IW•I W 1w . •latively high-spead flight with uch l•ed slnlg L4d yaw4ng foments amd to wind up the sinl. los to th fusp.." fwur-point wptiamon l•y. wi-.h fews ho AWIP witlk a WhhuS popoIOUI out, however, has a numbor of inherent problun So be au*.. te ck anOw di wg area, some of which ae uiq an some of which mams oh wv adi~u~a tis li~w~nup ariue. are shared Wo a Mow degre by the two-point fuspenSymine•wd high-denoty beA =mbe allowed to s.. rmuet Mhl retAMba It *Ok Minor eecta on baelEMePrgn relaw of mulipoint suspemioms wapkw MOi qualies. froielos mt ha a& for quir muilahneous jettison of &H cables with a hit ft roWaw m a Iil mW be J letwm the dme r .ofvability. If olimu ai ioworpaumd at the book end On pIi to pSrowdo Ipo rom swspe•mion points, theat hoists must be syn1Mm d• t ~in t q. gchronind. The pnmme of two or more autcimat Tme twoefuim iwnpsein PrWiAs t6) yaw MAd po while prowiding stability to the caWl. proPi"c -1-111'9y wIVAh ShROOIhimt VmAW=S h vide a load putb through wbich potentiall mucn-. trdable moments could be apIed to th helicpter. 7he dirtonalW (Yew) rMckesg mamma Nl of a twoou omispue a, as C faiiof WW ciisB lenth, is Honking ap multipoint wuspnsioms to a hovering lioptr cun i difnb ut. The operaton is dea-
-of
weow because of the pmasibity oif pickinp up a par-
N
7 L
R-L/do
(13-1)
tiniy unhooked load andsending the k~ioptr out uof control. The problem of hinteruissoe uraitacr Ii
+'
.,
performnance is unique to the four-point mupieo. PMo v analysis of this redundant structure requires
coa~iideration of both payload and !adhopwe strctnt.! stiffame &A wall as Mug membo ulasticity (Ref. 10 the practical application of a eer-peowt scope.skin. it has am found nesawy io incorporate a load ~trim-Iasg s)y@am to equblizs forms in the four cables
*
Without such a systas. flaing oftthe fehags and the payload in Pight is likely to load soem cables while cow*~ others to So dlack. An aiawmatic takeup donvice dasigned to allevimt th" problem is likely to be complex and =Lpensive but may be a necessity in
fliture helicovtim Uusif four-pont supnin.ML To determine the location and copeotty of the usaPomns points for amy new belicnowta the designer must have some knowkledg of the type of leads to be carried. Refs. 9. 10, ane 12 list, among item, most of the Army equpuwnat sand vabides which can be slingloided. Weights said dimensioas arn provided. Army aicraft axe hincuded in the listed beak because of the
*freiquent
uwe of helicoptenfrs faer i! recovey of air fore hua. io"
crf fro otewe bWu *
*
13.4.2.! Lads SSI* Tba reatd capaciy of the eawmrna load sinpeauwrm system is established by preliminary desep and flight tesb. ANl ompouvaets of the uslipesion tysteni should be designed uniformly for this rated load. The rated load mums be mulliplied by a limit fliht load factor of 2.5, with a sfety factor of1.5 appliedto givot an ultimate'diesign load fs.or of 3.75. The at-
tacbingmnactur muot be capabk of sustaining this ~tension l9Moa pplied in any direction within 30 dog of ventkcS. Relie obta-ifed Lf i
*s'uima f in Pigý 133-. TbIG deaAAWle dim~as maebsaw sum to sormm-etw fn dluhl mwmn bmifg miuw both
the eheliopier mimi dho - -g
ame emi-Id sa r
The natual fraqamemei of this unia~ I
+ ML)-
______
is
,
&Q wher K - qiria rate of suepeasom, h/ft UK- sam of helicopter. sft - as *(ofitral Wend. dug
Experinace his shom that, whim thisnataral
frequency reaches does proalmy to the Ip aiim rotor firequescy vertca boomn will o r. Th lowar practical throshold of vvgticl bournc bm basi defined as approximately 0.6 timeas the IP frequeny (Raf 9). With pruimndy used sing meters amd yia
load weighs, vtd*Wa bermnge km kasi gnaemu rd is dqrew rsaqa fronmO ld c dt An- 61 1or to p.tin~tadally destuctive eiverest mitm "spoes The midor cagu uuelly mo.th.- rosk ieco-bh lower threshold freiveiney with the conbisat ofm a stiff Flifg and a roitvel lgtk load. Mie mmr kmia divergmnt cour whin the wiliral fuequmeay is taken too dlose to the IP bWAnqwy with a siewwded load weigh*n clams to 014 impty wegbt of the be&co095W. -esi
IP FORCING FUNCTION
om the V0&S requiriment can he 1 eshown that, witht Lh@fiitflif
safety facto! of 1.5, applied at the extreme aftward tutag angle, as establised by prelminay design. Maximum towing force wil he determined by pkth & angle an? ucotrol powe lisits, wham the location ofK
the cohls atzachmat point is known.
SUSPENSION
LOAD ISOLATQR
DAMFNNG
13-.2. Dymui ~STIFFNESS A plewmmoonom known as vertical boumc can occur wbm a hlikopter is carrying as extersal load vAt-
panded hresa dliegor a peodant. Ittisadivapm.;u vertical asileflt"e of the airfram'/aarg system comile by memsortanc of the coapled airframe/carg UnQurl frequacy with the em-par-rev (IPF) vibnation frequency of the helicopter. ambis Ashw.
M
A
F11en 13-9. Hallesnplm/Leai Dymmseles S~kmaak
On theassumption that helicopter ompty weight These nonlinear load isolators can be based an sir and main rotor speed msuinvariable. and that the springs or liquid spriaugs, although studio shave helicopteir is capa"l of carrying its rated payload ex* shown hydropneumatic load isolator. to be geweally ternafly, the only way that divergent vertical ossuperior (Ref. 11). Such ao isolator usually is docillatlom cam be avoMde is by controlling the stiffnma signed as an air-oil cylinder, with a volume of cornK to keep the siaspeumson system fiequency, below pressed air providing the spring rate, and the oal proboth the IP frequency and the f~askg first bending viding damping and aLo a momn of varying the sk mode. volume to change stiffness. When the applied loed is If the characteristics of individual cargo loads are increased, an increased quantity of hydraulic fluid is known. it is possible to controN the stiffness K by metered into the cylinder by a servo valve. ProtIl~orin the dling &sign to 2;:lo; W by manipuvisions should be incorporated to absorb the recoi leftla Wing geometry (length). material, &Adcross shock resulting from sat inflight load reAse The: seadlma area This has beew done for some existing load limiter also provides a convenient mount for the helicoptee load combinations by pireparing cuve of placment of a cargo-weighing loud cell. minimuam sling length versus lotd for vanousamsupumsion gmimstrims andsing webbing thicknesses 13-."± Wkarhe mi HehWi (Ref. l3jý Where suspenslan length is variable in Whl mabebetocdu samopertions saytbefacbrly wth acdc W ciit load thesrna M&h by the use of a wfsickhe stfbe of~ mutisfe~ slingh an cargo tiedna looandoperantiors on be deigr, asd by reeling out moire cable whenever utlgdn n ag fm okadpnato the omsia of vulica bounce is detected. of hoisting its the capability have should helicopt.ur diof problem the to solution isausefacory A more vergat uicalossllauo isavaiable~ ~maxunau.n rate externa load while in a hover. Thi capabuiliy maksk. it possible to retrieve km&i frcim i toconrolthestifsigner(Ra. sqeebissoltin (Rd 1). 14)Thissoluion tO ontrl ~tight "ntwhere dose havaring would be unsafe, and amb iefac i 2 tthe zapnw yur a Iced ~i Nol~t toswhic lia ofe pdtoFendulumw malse tekndbe walatorhavinaga soft Mon"g rate takig advantage of hc sisin edlmla ido h mz in connected arm iprop two when that, fam fhe swss, th total spring rVat K of this cauhuat~aion 0 theae cntirobaility (Ref. 15).--.The detail design of winches has many unique zsalways less, th&A the stffens of SONhe wmig, in th pIOP~b~npocts which ame bayond the realma of thn helicopteir designer. However, the designer should be familiar the goieriti characteristics req"ie so that inKIK2with K -,P/ft (13.4) talli~ent speiificatioas may be wriuma. Military, he&iI+2 K + K2 cr ersued for cmistansportmasthbave aighK2 speed win* capable of hoaisin fth maximum exWroa loa ata rawe of at Ieat 60 fpm. With lighter .
WWu
locuim. SPO6dS ii*- 61MA ON 'AM 2
Ki - spingrafJM bitsPringh/It X2 - -m risere of asoond sprW&g 6/ft Thesw-i thevlod isoltor by itissIf hass a situffinso such thea th eaternas load foawqc is always belw the 0.6 X I? tlofhodW hMrequsc. rFuMý With IP fire*WM7 sid coupling with 1km fumehrge besidug moealways wil be avooded, rqardlu of ths seasp -mwk a sing ssfuses, iffa dowle spring is empoYed as a load nisaie, she sam Mile (etltsisa) mVAs bi MRmoum* ot to bandi doe ighees Lend euhhed with dhe Masti almasi. A spR&n4Mp Weeate deega so mWthes rpuare. him Probaby WE eusmu Mn irdeio 4140011when11 loade with ase meXWOMr cap land (times twhumi ked faer). YU mone isapsinsbcsc type od loaid Mom ean we sin I ho ane a vow*bh gmigs" that
A no-loa tepkoymes rate of 300 fpw or more would he useful in a combat situatio. The wia speed siould be inmmothiy variabl over its entre ramp urnder load. Adequate brakiing cawacity must lie providd, as wfal as an automati loal holder to Ilock the wiir-l in cowe of Power failume Hydraulic power smsrally a used fw high-capecisy aircaf weacha becoame aircraft ekutrical mot~ors have ineufficisat power. Loth drum anad capsita. wichsm wre ssitabl for hsalicapter applicatiom. Stadie hav shwna s dewr advas for eother type, when all spsets of thmi applicistio are coomidmd (Rs. 10. 11. ami
wINIa muiat ouuiM Asupeasl qus~.newts
rupuc to smenemum alowabhe ik *
wma matena fho.
are
.PWY
.:LL.bkr
16).
The same etUdiie have cancluded useenum*e thatwith the pessiu ftet 1 su ierp a tUN optimum teasms member. Cable minuw Mdeave i die (400 Isiess wrse srand diesimar eran everi
f.
ofd os Mk diameter). Mmoabu roiW~s rooc. 06 the other -WW wuk*m Iceh M.COMMeerkva MM&d med 68 4am dmuld tamblih high allow46 budI by IN it PeI We CoGmatau.peay elecbiso e, A quire a number of electriW*~ tipw a #ad citolom so be bid w"th the cam of the ware so%, to be wed for powerigi Cis book-*peatng aak It uadoid eSa Gmdauctin varm iondiao uIb~ bevera Incwrorte the designer thid ke eiMd Wum *4*0 withi the cabl so tha Chug quar wire mbe
* * *
*
-,Moad
a
usm~imes
for coAnduro
passenger pod opw.ihions mumt hav amgne GdpOuitiwly kKlokn the pod in phe= afte heiat or of piecing the emergonicy jouihon syuiec to apositivel ukf configuratio, Safey Of ron-adagpersocael shu~db whaiscd by providing some mesa of etammavedw static electrioity disebigs The tAntic doctiwal char developed by a arm"hefiaopote hovodringIdAy sed duuty conditions baa enough aingy to Weepcitew an anmivldu .0osilgin contac wit lb. aMopues book.
This soa
o
arigt
ignite fued vapo or initials explosives (1Wf. a7) Twc hmkma deuimg ve. techniques wre availble for the edimation of sa"i ~~~CKV books musW be capabl of rebesing by elseChfirges actve discharg (Re(. 18). which Mo ahisktuIqel ciommaid. ad aslo sboq~ki be ask3 to be re"oktag &OdMeto tc null any Whelopler poteiaW 416Am incal u system. A ftua IGAmd bY a bw*ckup tocord by an oobad sme.or and possive diinckbar groma o at" reWon featuce. which oppos the book which drains the capacitive enWg to grand thrtwgb abo is desirabl. Oe singliewhim the lad.iskasedl rski iAk,at a curentl Wh ichmis a highly pdgtsospaswkwAaweiv nmusbe pro~vsd to imuWjusiu and usdtua nwt dctuuble. 7he nuiaii' fte the hoist cable from load rotation. If an electrilink can be incorpore in a posead handler' gaff caW opuwad book is wsed. elewrca coetiauiq mugt bouk, or cmn be coonnete in swam wit the sm-. be mWAitalmed through t"i swivel joint, using welldiinaing. $so pbr. 7-9.3 fkw additional informsnmWndsip uingi. ti. The r~fhummts for winches warhooks genrally aM ulmdw - wheth erlsinl-, two-, or fouar-point v~aa~s arO - a-N. Heck: and ez-.ckir for mdoaempuosdo Not require & ivel noW- 1 .P a n .W kWP*~w hkstv. Four-poWn suoso book mus aa I.vH.hf P. VaFotasndi. m W.o ? &lua dwce Do, ~ E~.~gFw omul-ims aukm=@* aIs-i- . unwnwder ulxims. Whxncham '&MW Of Wfqku Systems. WADCMh W6 t~wlet wil w-m'Lim nowt be syncbrniaviJ, cithci /M/ Wright Air Davslopesww! C=e, Wrigh488, a tbg tkwosqb or ri~3IaM8*vO~ do"t amiasdcal Air ForeDo@% OH, October 19%4 (P§ Patterson A! WiROAci Mlduiiphs uvols. esr~ alaor WM of an@ w ~ to ermit w.a 2 121916). siculd be ldmdiiunl w*Uirl" N U 2. Wnsk Peed Ceatw#0*I* tie* a * an ~um~'4 mecuq~-Po,3 b~iua ad t0 O~Vid a 61-91(C). Aircraift Sysesm D~ivi"r Wrok. means of-trimming~ tree towipcnt load in nlight. Patlereon Air Force 3... OM (AD 2MM4 (C)) 3. CrMA SWW Dvcs Goid* ISAAMRDL Th "tfault fiva #Asdly; uboeM be Emad early ill the daiga of "k wdir.svt carg syeaw to isolae p0bkWat haWW~ CANIal k"aWe must ha ProVi5a 00 #A casW9aJ suspmeom sytms. The wrmal release 0m49*1, wlbIik may bi. elucarial, ame be lcate onj fr* plIIJM cydi s&iU. In addition. an allmhasad~ mom~ of Iced rulsan sould be pro. uIsi go a tooik" to ans axmal Missm deviaL An ninsj q sary raslavmotol nawt bt providedl for he
~PWst This vat*h ensins hema fth unwu and M-'
.mosietvi
mW& a
s ibafa"Sk"ou
jar.-
biaws of the apik and hook. a~be by soamgq the =%*u~) or ky veal dv&hmh~iq die winch &=m w doet *a bdg laud Ivrap fb c~w Weo the ddun. Tw*-P*Ua sonassas shoul iimiepon-ar. a &Mmumue whic deassie the psotinfa of ams c"~ "adhmu sam"""Who" tae Mmrvlvh cabl. Im-1 OMadaas.
in ONOM to be maa-e4nte for
-w.issfil
j
tober 1971. Sa*d. HEL M24-73 US Army Humus Esiginveing Labors.tory. Abordeen Ptovirg Ground MD. 3. 0. Koaroukls, J.3J. Glamy, sed S. P. Deer*dide% M~ 0=61%. D*Wd*MM. MWd TAMt~ dV
-. Swoaedlxed Cb&CIM? Cowra.
Aircwa
Rua#ulm Spat.
for Amqr
Abwo,
USAAPERDLTR M276, Emil DiENmmrms. Farx Eusb*e VA. Jua 1972 6. hPftrei 1w A#mdo wwdA 4Cv SaSpprni (Laud Lbj ed) T.ev Ser for WATAS. USAUWAL Memon Rqper 72-5, U Army A~uscy ior Aviatiom Sul"t, Fort Rusher AL. 7. W. E. Iheauh, 31edp Gh"* Lad Amo Simt pdwe. 310 ad Abtv"t AfJd Npab. EdDkemovir, Fort tSAAMRDL.TR 72-1 Evo, Eus~tis. VA,
JMy 1972
I
bi4ioArmHo" IQ 5dsLow. OW Kw*n. MrVAwwyK. H Wtt W.
.HbeLC etaDW Fo~
wnadAnlbo H u fteAm,23
Aviation h4Rv"ka Cowmnaic, St. Leeis, MO,
HavtL~fA HMkArA
£uirind Carr kzhklw%
I.. f.i~i
naacS
4_,%Uisi aJ. #!aC tf.4%tbc
5.G
St& 10. L IL.OmtMO and H. B.tabin, Dm~ La dRA xd W L4*) Wiav Jfapw by* SOW WAAYLADS TI 67-46. US Army Awklith Menoial LIwAbrawlu., Fort LAis,
VA, Novhebo 1967.
II. T. Lammatk.
.d R. Kakipa, NRmv-LOh Desie Study. Spes Hsadhut Cat!.~x
TR 67-51'. US Anny Aviais Bak, V, OObW Musk LamuluimPou bllaia Lauseda Fat Bal VA U&14AAVA
45 inswy LQ# ft drew. Swta. 1Ec. I~i0~4M
J
Ammvpt~ag
. Wikx ~ LJ.
Askopt
A-
fuiatyi. 1t 1ý,17
,%1e
-".VDww
99
vainM~WLrrd~,Fr
1.09r.4,a.
TmE~tkCaf
ieagn
knw.
n
fHlcpe*i
rFsbrbi.InAL--629,Pr1. A~
~c~qa
ns
knk
I7t
ound yami ArFn Unt Wright-Puztwaroo Aflb OH, May 1%9. nieFecawlc
ma&hsawado
USArmy Aviation Mataril I~ratoria, Fort
Pua"M VA.
WS
Naugflr Etwa'oi $iaw Sp~a. Noth M&,shtAn. USAEMA ?tepnr US A-my 'hAWApoaimt ES*Jwamuh 4 A'riw
x"4
Ajq 1&
ry
wkVAdI@
&~lI
Al~fI
f.U~l
Coi4nc'i~oM ~Gr~lkkt Vo 37. taltumoic: WMWT-dy F'vst. 3ne.
2W. ASYfl'E, Hniql~.
34.r6jbj&, 14.G. K.~
4'0-14,
"'e
EMA No, 6,
U
9 21. AFFDL-TR-70-174, Awlysis &Ij9htninIwwors igh: Oizpkeyý Syir.ýeu. Dow fior El amct Flk PFK
H443
9
AMCP ?W062
CHAPTER 14
ARMOR, ARMAMENT, AND PROTECTIVE SUDSYSTEMS DESIGN 144 LIST OF SYMBOLS
aack*helicopter the ability to survive a bostile as-
A, Av A
-presented arm -vulneble am singly vulnerable area of component I of a
A V4
- siq0y u
am of a
INTRODUCTION
As described in Chapter 2, AMCP 706-201, and
"AIS •I1. the concept of bepter survivability is
-.
)
',
41
I
,:
...
:•.F,
broad aid complex. Howevr th effects of wrmamaw, and pricti cunthr~ can be evaluateo i' Vterms of three baimeamo of coping with a&tue envruonnent ( vbeber the hosl i man-made or natural): I. Avoiding it 2 NeutidWng it 3. Enduring it. Avoidmc places emphasis upon vedki performWmc. To avoid a specific enronmet, a belicopter requirce the ability to fly above it, or aound it. or to outrun it. Armor and atmament have a negative effuct upon this aspect of survivability, becaue thgy add weight and aerodynamic drag. From amilitary sandpoint, neutralizing th esovirounnot reqm destroying the sour= of hotlity or hazard, or otherwise ren6tiag it haslesm. This may be acoompihbed by Natie or tati that do not involve the system being pmitecaum. However, a capabity to correct a hostile emvirommet may be designed into the system itself by the inchnlion of standoff ,eapons (aram=et) or counsermem us. ndrias th@ bombu advirowmt plum emngmW Endui eupo such arom a redundancy and eparartion, eeing of M dom-critici components with strucUTar or componensw that amres "Ascitcal,concentration of citical components in a siagli plam within ft vc
c
in orC t
rma de-offs among performance, armor, an arm- rt always aff=ct virvhvety, but Aho may involve m
dimVly s-o
PkIH - conditional kill probability (P? IH),- Conditional kill priabty of component or a system
14-1
vironmsnt, although ematl for mimsio completion. is not in itself a true memsme of the worth of
the Uo" am
presentsd to the hazard, and additio of deadweigt vfole" on madb 0 amior. Armor and armament, eampeciafl the latter. must be evaluated in conutres ot than srvivability. For
tube pramary mission v
s pe-
forunoce (ae Catpter 2, AMCP 706-201). Te major tatdo-off will have been pefomed pror to the start of dtal desin. Altongh thuse ttds-offs should be updated contoauously throuhout the design procu es more complaie and pr•cie data becowm availabe a diceggoei of their natRue m d ript is beyoud the weo. of this dapter. However, it is weential that the deinr rinoba co. e t t belmnowproc to veroau mmia 0 detMail e decidow regarding location. intamUs. tim selection, and an of armamenL Theme rninaioships are subtle and, in many iacs, arem dficet to quantify. As a resut much of the mewial oamtaod in this chapte is qualta•iv• with s ific m latiouship being included only in inefaems in wbicb gmal physical rdationships am known.
142
RMAM
T SY
Helicoptier armament sysms typically we individual installations of guns. guiWdedsmmi, or 64eW flight rockets. The namber and type of instllatiors, and the mix of guns, mimlk, or rockets, we dmcsibed by the governing belicoptder detail pe& cation. This chaptr contains the bu*•c helicopum deeig guidelines, comideioand requimust that apply to the installation of armament sytms. 14"1 GUNS Guns of the number and calibar decribed by the heicopr detait lt stalled in the requir poston. An in the rquanitey poutyp ictate A m0Pply ofthe quanty nd type dia dog be provided. c poe rp ktio
ad be eammunition atheutio y by the dail
14-4.1.1 Typos Several type of pas are available for bhicpt usc. They include air-cooled, -opabd 7.62 wo 14-1
weapons; air-cooled, automatic 20 mm weapons, aircooled clecuically-operated and controlled 30 mm 40 ien grenade ammnmwstion- Typical guiis are listed in Table 14-1, Operational and maintenance detaih mvr contained in th. applicable 6pufifcation. US Army Armanient Command (ARMCOM), Rock Island, Ill.. can provide information about g'wns not listed in Table 14-1 and also can provkd, operation temperatures, 'ubrication requirements. power requirements, peak and stadly-istats recoil forces, life of gun and critical parts and similar pertinent details con all Sums availabl for helicopter use. For -smmI applictijo s, the gun is installed on t=Wi adapteer Thia installation minimizes the effect of gws recoil force upon the helicpte structure. In sislcting and designing recoil adapters, careful conWieretlon alsuld bo given to gun muzzle: energ; gun weight, including attaching feeder and drive motor, and the respons dynamics of the helicopter structure. ARMCOM can provide thw latest information on the'ss requiremmat Feed mechanisms for thes weapons vary, and eswh tyepresents a special set of design problems. in ome Instances, fte ammu4Ition Feed mustat _rt from rat and must reach the peak rate cfrar within the tims rcquired for the f"in of one round. This imposes hWg acceleration forces upon fte feed tran and the storage container. In other applications, a feed mechanism is used that extracts e~arMwtr s from a recycling conveyor balt. The integral feeder, is adapted for this purpose by replacin the feeder cover with specal link guides. 71a high acceleratmo foaece imposed upon tie belt, j anid cydkW. devastioss of the gun; from a nominal raes of fire, require ceful design of the ammunition
food train to insure equiValeAL belt tension on both aides of the Sun feeder. tive accelerations upon the ammunition belt by requiring the belt to start. miove one cartridge pitch distance, then ruturn to rest in the proems of firing each shot. Sharp turns in thc vicinity of the feeder, and belt drag conditions that will cause link stretch under these conditions, must be avoided. In some systems an ammunition booster will need to be added at an appropriate location in the ammunition path v' that the belt pull forces are alleviated. The booste commonly is driven by a fractional HP motor and must be controlled by various means to sens ammunition demand. Starting and stopping accelerations must be controlled so that feeding will be compatible with gun demand for ammunition. Care must be taken to prevent obstructions from falling between flexible chute elements. 14-2.1.2 Locatlew Gun location is part of the overall heclicopter optimization as discussed in Chapter 2, AMCP 706-20)1. Thlere are minimumn requirnessits imposed by the naturv of the Wasawl. TUC location mint Provideaccessibility, unimpeded projectilic Moit paths and debris ejection paths, and the ability to jettison externally -nouniad gun pods. and it must be such Juit fte vehicle can with~stand gun niuue bMast effects. Beyond these minimum requiremuents, the degre of optimization must be related to the overall efrectivne criteria used in avaluatng the vehicl. The: trade-oft. will involve structural and gasmetrical limitations arisin from the desirc to optimie osw and fliakt perfnrnmaq vW qh@woaniniock" a
TABLE 14-1. TYPICAL. HEUCUIPTER GUNS GUN DESIGNATION MODC MACHINE GUN ________
M134 AUTOMATIC GLIN
CAUSER
RATE OF FIE
METHOD OF
Ofi1
OPERATION
Wo
SELF-POWERD GAS
7.62 mm _____
7.62 mm
EXTERNAL MOTOR
___________~~
MselAUTOMATIC GUN4
2D m
M127 AUTOMATIC GU.N XI40 AUTOMATIC
20 wn 30
CAMP"_____
XMIX AUTOMATIC G114ADE LAUNICHR
142
4(10-720 4D0-19D0 VS40 ______
40 mm
400
R~EEECE
RI~NG
ELECTRIC
TM 9-1090-201-12
SCLEENID
________OPERATED
7510-4M
EI0()
PERCUISSION TV 9.I0O6.AS-15 1-CONTAINED)____ SELF___
EXTEA.... MD1iVR
ELEC TRIC
To 11U01-IZ-432
EXTERN4AL MOTOR SELF-CnvJTAIN~ED
ELECTRVIC PCRCUSSON
NAVAIR 114MI97-1 POW 1005.~-41W5
ELECTRIC MOTOR
(SEL F-ACTUATED)
EXTMf~AL. MO1TOR
PERCUSSION (SELF-CONTAINED)
_________
TM 1030-2013-12
)
All" 706=
to provide weapon system perfonnanoc and maintunability. Paramers relating weapon system performance to location are diacuzied individualy in the paragraphs that follow. 14-2.1.2.1 Prsec!1 FIight Path Gun location must be such ats to avoid intersection of the extremas of the projectile flight path envelope with the helicopter structure, including the main rotor and externally-carried atore. The projectile flash• path efrlope io described by e circular diap-er "sionof the flT4d proj.'til with the circle center being coincidentinwith the gum barrel =enterline.envelope Factors into *--ader determining the dispersio dude gun and ammunition dispersion, saerodynamic force acting upon the projectile, and defaelions of the gun mount and helicopter structure. ARMCOM can proride details of gun and ammunition disper-
J
ý ,Juzu %SU
For some application,, an ammunition feed ys•em is used that. retasas gun debris to an internal stcer•a compasrtmawt. Thi! design is suited beat for syst• m
fck
Vm LI&
_t..r__ _ 1.. LM aAkAIWU ma 11" GO &A U.
iouci.
anism.
sion, and aeodynamic test data. 14.1.2.2
rotors, or externally-carried equipments. 7U trajectories o. the ejocted debris can be determined from gun ejection velocities and the local aerodynamic conditions about the helicopter. In gemeal. the dAIs ejetion velocity i3 equivalent to or higher than artmunition feed velocity. Debris ejection veloities can be inciased by the nuse of accelerator mechanisms. Some accelerators ue rttn rse ocpueadkclrt h ers rotating brushes to captureeand ctionlote the debr-g othe use sprocke. The slection of the desg tas nique must consider the available spect, and the at. titude, kinamatics, and shape of the ejected debris. The accelerator must be designed for positive cepwn and retention or rejection of the debris, and must be Iato as doe oetto the h gun u ejection jcinpr port as is l pouls. tke Deflector plates can be placed strategically to redir" h i ejetion pat properly, particulrl where cam can strike a surface pipendicular to th
.. S..M
HM
L-1 UdA-
copemr strucaure to minimize the efects of muzzle blast. The aircraft skin near the mtriule and adjacent aimat aftructure must be strong enogh to prevent gun blast dama. Reinforcement requirements of the airaft skin and strucr an determined by the distance betwee gun and skin. the thick.es of the
ski and th density of frame and Sinngm. The gun
that otwwnlcv tem..vinm rna,powgnu nr anmtAkeda wathw --
-""-3-j
-
~
--
-
-r
than a linked belt, to tmn ;port ammuaitioa. Tie return conveyance path sheIW acoWmmdate positiv guidance of an oCASional misfired cartriOd as wal as fired cams. The strength of the storpe comportment must be sufficient to protot persouu and the vehicle from hazards caused by nmisftd crllrhis. A
ventilating system in the storage wntainia &W be
snmuz never be locuted near erough to canoprovided to remove residual gamsin the fired ems. pip, radar antnnans, or door frames to came or cremte a hazardous condition. 14-. .4 Extmud Gum Jettluieag For soe gus•s a muzzle brake can be incorpoIn general, for installationt requiring jettisonthe ratd to reap the blur pressure find. This evice Igun and ammunimion sould be located in ex•t•aydistorts the blast field so that peak pressure end im. monted pods. The klcations o f the pen and pod mad pulse aev rotet. and displaced from their normal On an& of ejaction salu be such ea to insar poastione relative to the gm barrel and thereby dearance from the helicopter, landing I1pr d ad. ducs recail forae. ARMCOM should be consulted jacont stores. Pod and helicopter structuir sW be for details regarding availability of muzle r•e s d for compatibility with power jettisowaig Ie and ft" iMPPrMM for Particular weapons, dein MI.-A-8591. wAdentiom for fiting a particular imn. and definition of muzzk premsur fields. AMCP 7011-23l pro*4*.. Aej ty ff provide Locon of the un mt tO u a vde information on the dig and
muizzle brakes. Consideration also shoul be given to
the rdeltlowI of diffrent we•apo systems to each other, ock as machine gum vs rockets or missiles, bot, fro beth tr s msstitiomm yy Ne fiin positionet position (of machissies m achine gPa) and during the trojectory (of rockets, ctc.).
"!144.1.
Deke Theein Paih
4smad mmi ps
cam sad links &W ow mm
puporn heicoper structure, cotrol surfac,
cleAn~alnd afctessgbilitytolwperovnid
es
offi
lence and acoibifity to &raw pefonce Of maintenance - including swvicing. removal, mad i. rsaml lo in mt a dofuthe l agua i g and ofa related u itoacasusres; . T e w saMdm The ammunition. of loading ard unloading also must be accesrible enough to permit suc mcW tivitis as diagnosis of malfuftiom and the Ari-g of stoppages or jams. parial disasembly while in
plaow p•formance of all andard bdusteatts with appropriWte tools, and viewing an, reading of ad 14-3
diask and gass. An unobtrutmAW view from the gun barrel must be provided for boresight .ligimnML Electrical cabler, tubing. 54 Oquipamit that must be placed in the v~it fthe gunJ abould be k)taud to) preven damag during die removal and replacement of the V=n Mintimum maintenance and turnaround times Ire Coindeutiis.Guns ~ti3Ik
Cables and eqipimwat should be restrained sn that they coano be proximate to hot gun barel or other omoing parts of a weapon or systemf. Provisions should bns mae& for reoving hot weapons (gun barrels and &4oiaing compofhuts) without buirning creweata Apropiat bn~ls "NOSTE" mrk.
inp will be Provided on fte weapon turret, fee
ehuting aimauitin ague, and other matenel not du~m towithtan seapen foresshaped
4
ý7dewhpVMeut
14.2.1.3 Types of Isidaladoes Guns n~ormally are installed in the fixed, forward position in pods, or are mounted flexibly in remot"controlled turrets or upon manuuily-operated pintles. also may be installed in the fixed, forward position without pods. The corifiguration that best suits the detail helicopter mission requirements should be determinad and employed. 14.2.1.3.1 pod sintltdoa.. Gun pod installations usually contain the gun, the amMunition storage and feed system, and thc
operating mechanisms within an aerodynamicallyenclosure. Size And contour of the pod trnvelope are selected for minimum aerodynamic drag. Construction and suspension features of the pod shall 144.1.26 Dytlcý Forces conform to MIL-A-8591 structurai design standards Dynentic forcesi may best be determined from infor jettisonable stores. Pods normally are designed a on turret ww.pon actual of the taut strusnated for installation on either 14- or 30-in. spaced bomb miount that simulates the flexibility profile of the helirack mounting hooks. The design of the supporting copteer. Weapon systems contain a series of shock albstructure also should consider: .. eb- tait d!parta w~fl atual ~icea, 1.Location of the pod so as to permit normal serthe during profile load the changng thus afteruatiy vicing and maintenance of the gun, ammunition, and fn"i burst. Theas shO~k absorer.s am characteri-zed operating mechanisms without removal of the pod. as Theme actions will include, as a minimum, amnmunihaumisi 1. Weapon intairma 2. lafie ydralicpaciep)tion (sringor loading and unloading, boresighting, component checkout, and normal removal/replacement of 3. Tesre stIauCI Components. 4. Coanneclin struts between turret and helicop2. Design of the supporting structure so as to withter 5.rameork llicotar r ~stand forces imposed by gun recoil and aerodynamic Simsthirewepoc atvarius ttiudesof lepressures. The structure should provide proper rigidity in order to minimize gun firing errors as a result of s van -The torasandn . . nd a uLbth vau strutrldfeto. bw inod ntofl is satic~y indeterminate and 3. Location and design of the pod to avoid aeroShoul bit obtaimoed hra ARMCOM in the form of dynamic interference amo g control s'urfaces. seninimmeanl firing data wit&halost cond iooi clar and adjacent stores. a w~ntifed lyto:sors. 4. Asymmetrical firings occurring due to failure of 1.Aismuanhion Lot No.
2. Wespon Wei 3. Firing scbxb*l
4
of the shock absorbing eleimets in the weapon system.
4. Wespomn dirasiom/aznmwkt an&l for each test es~uemmTurret-mounted S. Slilffum of mount bue etc.). 6.Othqer IPaltesconditicab(temper Weu, Modemis ovea ui in.the weapon systms niounsang iinterface sbo'Ad be considered in new or I is to allow for normal -growth" or Ineem is. impuls as an evolutionary factor in a again Mitcy~clic A.(Ke s of 50% is recommended, AwdummAwW*. this consideration will provide an in. --i'nad ineis eofapabiliy toaccommodate unpredktabl advierse conditions, or necuseary stiffening 14.4
a matching gun pod.I
14-2.1.3.2 Turret Installations guns are aimed and positioned remotely by means of a sight or a fire control system. Such guns normally require remote location or ammunition supplies, which are connected to the gun by means of flexzible or rigid chutes that guide the ammunition. Turrets normally are placed just beyond the helicopter skin line, and require an aerodynamic fairing in order to reduce aerodynamic drag. The fairing design must provide clearance for the turret internal mechanismit, and must minimize aerodynamic torques against the turret drives. The fairing,
V~3 WMCaa dMMiUi Owe "BOW" cOidizio Of the w=npoQ a.4/Q*g tube w4"Okm amliua to tinier the Iipo
sae
wetin urre AmUperik us 71W oftht
4
IrO2t
K~~
Al
M_____ Commlhs of hM fiki bond is evidw *A to p41 me of the inu Fle OW. e dM bOW do tWw OmwofSatopO, dMAb jassa w pumtW &GWMtd inm~ataLs If kq is nt ch,
mM ~emitew a~mr tileW fijetry wtt dou#Wu Owc enugr tearft o tile CG/Oecty &%K do*=es d ieanu il ueuicwe nt q s prro- blema. W% to bpca pwdmd bwom tinoiu bapm. a infti inflo ms S pdhum9thu*~a Vs. uS&M vTcd w.p ordp~er to pmv'n tbauumatkme tuuvinM boyordpotheko rwaiabs v4tota pwer.a aa uponx toops ame bhepgum med ioun Oftjo W uw asea i tile Uv)OcorYo whgj" &do tcot funors. awat QUitkUrtr kh a omfwtnW obdi.iots wiartiis dleard m" bsdfratchetgwe prowiked am wVil oý& thepowc ueic toathek to PA tetu tw oino i het *.aeekAfj~w fuoc* th potbila.defiaite toWreationhupton pasr heleteP to nds"OOOOS toaWcapolt~ia- vatfoOle. ( set factrs. Quick-6) The p of..1 AugnMA, A V af W**aonntSouphapDtb-uW bea usW oasuracesmor HAmmNito Whi pmbe &tlu eor~mw oef te "C2c plane the thelopetu hc l rw.cýM.smsus~#bp~uhjulo I hat.i qiwts must pirconu estb a~ n efer-a~.sefcto.Amsle otimmMb
ivtcallozpad wie thes fromap theflo Wigal conduons applied ihtsiuad Tay o~nuouniwsoan: I. *.uPal~ ldsU imuttt probyite tancureringioi hehcl~aiccpMt erC'1" hCyci~sndth.inechaiml fmpose &dby m.no~x theW n. byth u a n h rni~ ký ic. lorquswill uiet from ar oloins erth tondiitns, mPWnthrdvam veoitn or~lr unio ln:gt g~~1 to rest froau itmrkA briy thei 1a,vuoit~'urog 2. Eyriai :.,oils sach RhCW t bos Crithe hydisthe~ fun Vt NOSEP arif Acc-usi appliedo tW taoxua mustbe adeqe&an to its-
mArntu
luaaia veesaanye isetion oi ori tb#e pC)ivo
gun to testn fhomuld pemaxitw thelcityIu! a to-
froviu. ~as dsn opter$m b f dk saiikttioven. 2.ag" pmmcsheoooakiiW(Th11duiPw w aos clorsde t*othef oim as as~ dmapd MI hem. tain'te.) ae i-46 o k s" accetity 3.Andna urvelzrr nsbvw rW:re y te.fa coLdVAUnwed desgn Maiaaddu irietse tis-t nsu bly of fo:'h atcin e edai a qs to conualer 2i.tn mabaseemblimte.oiia nedo mst omjen Fof -p4icaonen, ofmuitiodontnrs haun con
3.Pri
on
wkila e
yf
ingqunickdinrdsion.ec icludlingkriaim.ed 7Amb u ofattaohind containes, aW hpiana uofl atyreuetrnrudting.
shotMne be consideed
Feortasomer shoulcdmcontain sufcmieacoem m th a-
4.[ttrnpth storeai eha&tosito.Te creiing byuk flntsuoj wn as e meurPoer appropiatel sou tbe j~au~Ws .qaur thro. th-, "N Sln ofiwil iugrnv. ousting fuickr'es. When usmplikedamunti Ahelicoptermustu ofr~io shohl bert ibt accountotr musefood prdoblem hmeben traners twhish asignfi beOnca#). andGinspeto wit mainenanci ~ypcaiy inhelcoper istsle. bor~y~ itleVibantioy rdcauetra auingoft'ns uprbed loninkred.mo gunseawithout rndquiring tho wuay.- toawnn muiton maig aer f odifcuttopl coniiner ocsI~ usallyarc raaoily n,.lariloitapat. Use of sactiashel i es lw or alops anoud fiomthe sthfl poibtion Theoegca. cowigsolnt nsidered; also notepropratel tnsor ticnklens wi P" hrinssa h wthe line of u! uwhiebingo crovd emontaic-, inate fthis roem.When usinge ofitmmenitomn. prov4~dtha wsg~tolth he ulli~ninuntbn cn heli cop ytwter efiunon hule takdet inoad.cournt nayps y dWi n chope'w.rwr* pitevome. Vibrtionca see adinC f t upg-tdlinkd a-5
4
AMW 706-2O2 1-2.I.5 Ammim.lon Feed The chutes that carry the ammunition from the container to the gun feed mechAnisms may be of either rigid or flexible design:, but should be fr-A of inSide surface projections and independent of other accessories. Ch,:•ts must be attached to the feed mechanism by a quickly detachable means. When locked in place, the chutes must remain in alignment with the ammanition boxes and the gun feed mechanism throughout all adjustments of the gun. Feed chute design must accommodate the aliowable twist, bend, and fan radii of the ammunition belt. The detail design shall provide accessibility to the amr~aunition belt within the chute to permit threading t'ie belt through the chute and inserting it into the gun feed mechanism. The feed chutes shall be of light weight, low friction, and long wearing materials and of a gage that will insure maintenance of original inside dimensions un, :r ordinary opcrativ:g conditions. If flexible chuting is used, an approved design must be employed. Where the aznmunition belt contacts the feed chute, the design shall provide relief so that the ammunition links will not drag on the chute. This can be acompliasW with tracks add%-,d to the chute beads rolled into the chute, or with a clearance slot cut from the chute co that the ammunition will be supported by the case and projectile and not by the links, 14-2.1.6 Boreslghtdng and Harmonization The armament installation shall be designed for compatibility with the rights and/or sighting station equipment dejcribed by the hcelicoptct system specification. A means shall be provided to boresight the gun to an accurate coincident relationship with the sight. For pivoting guns, a means muit be provided to check the gun pointing angles in refetrnce to the sight commeand anglts. The helicopter design must include a definite relationship between armament installrtions and the sihth; and fire control references. This is accomplished by establishment of a Weapon System Datum Plant;. The itzstallation and triverse of each weapon, and the sighting and Lre control equipment shall be referenced to this datum. Accuracy of the relationship between the re'erences shall be in accordance with the governing system specification. Gun mounts shall be adjustable. and shall be .apable of being locked in the transverse and vertical planes to provide for a minimum of -0.25-deggui' adjnstment in addition to any adjustment required to overcome aircraft manufacturing tolerarnces. The deWatdcaign shall provide for the use of standard borsigh?. telescopes for performing the boresight 14-6
operation with the guns in place.. For turrets, provisioioshai be madc for checking alignment of axes to the aircraft datum planes (vertical and horizontal) through the uwe of the standard boresight telescope, with the turret aligned to three azimuth angles as a minimum. 14-2.2 GUIDED MIWSILES Guided missile launchers and guidance control equipment of the number and type described by the governing helicopter system specification s&Wll be installed. Currently being used for helicopter applications is the TOW, a tube-launched, optically aimed, wire-guided missile. Deaiils regarding this missile are clasifie, and, with required Justification, are available f:om the US Army Missile Command (MICCM). This paragraph provides helicopter design standards that can be applied to the TOW weapon system or to any other miusile installation. 14-2.2.1 Locntcos of Laumnber Iu~llaelm The primary function of the launcher installation is to release the missile from the htelicopter without damaging either the missile components or the helicopter. The launch mechanism should be designed so that the missile flight path (during launch) will be directed to position the missile within (1) the capture envelope required for initiation of guidance by the gunner, or (2) the flight path limitations required for target acquisition and lock-on when tsing a homing missile. Helicopter missile launchers generally will be installed offset from the helicopter centerline on armament pylons or stub wings to protect the tail control surfaces and rotor system from possible immersion in ahe exhaust wake of the missile. Good design practices include location of the launcher on the helicopter to prevent: 1. Engine compressor stall or flameout as a result of exhaust gases entering the engine inutke ducts 2. Exhaust gas impingement upon, or ignition debris collision with, the airframe and. all rotor systems 3. Harmful corrosion effects as a result of deposits of missile exhaust residue within the ergine or upon other components that arv not accessib!e readily for prompt cleaning 4. Impairment of pilot's or gunners vision by flash during firing 5. Excessive acoustic noise in the crew compartment during firing 6. Pitting or coating of the canopy by exhaust gas and debris 7. Aerodynamic interference between launchers
-MCP
)
and control surfac-s, sensors, and adjacent -tores The dign and location of the launcher installation should be such as to minimize corrosive effects resulting from the exhauit particles inherent to solid propellant missiles. Prc,,per -insideration of nmventive or corrective methods, including cleansing of affected parts, can reduce significantly the possibility of structural cirrosion or surfac- damage caused by motor exhaust. In general, missile launchers should bc locaked as far as possible from other parts of the aircraft. 142.2.2 Structual Cearams Adequate structural clearance shall be provided tc prevent interference of the missile (including fins) with any part of the helicopter (including adjacent stores) during launch of the missile. A clearance cone of 3 deg half angle, measured from the missile longitudinal centcl-ine at the exit port, is an example. Definition of clearance should include consicderation of the aerodynamic forces acting upon the nissile at laurch. These forces can cause significant variations in the missile pitch and yaw motion, and in the iin,.r displacement, during the launch phase. Sufficient ground clearance shall be provided to prevent "launcherground contact during normal takeoffs and landings, and during hard landings at maximum gross weight. 14-2.2.3 Blast Protecic The heficopter designcr shall pr# -ide strength and/or surface protection for helicopter structure and exposed subsystems that is adequate to protect them from n,;ssile exhaust ,ffects. These effects include overpressure, heat, recoil or reaction loads, erosion, and corrosion resulting from normal repetitive firin•g. Details of these characteristics will be available in t-ei weapon specification
circuit tester, or a single-point elecrial quickdisconnect. in the indi,'idual missile contact cinrdt which is suitable .r me with an eVterna circuit tester. 14-2.2.6 Jefttlamng The launcher installation shn include provisions for jettisoning the unit from the aircraft turder all norm-, flight conditions, including undetected sdeslips. Launch structure shculd be designed for compatibility with power jettisoning per MIL-A4591. The anle at which the ;auncher is ejected AaU be selected to provide clearance with the airframe, anding gear, and adjacent stores. 14-2.2.7 Effects of Aircraft Mauemwus Structural design of the missile launcher installation shall consider the effects of loads imposed by maneuvers of both the missile and tht !urim-aft. 142.2.1 Types of Imtallatmion The launcher installation should provide for effective missile deployment in specified tactical situationE associated with a particular misile confvigr-rinn. Factors affecting selection of the launcher configuration are launcher size and weight, helicopter speed and altitude environment, and ground-handling and loading requirements. For helicopter applications, the launcher g.-erally vill be a fixed installation located on a wing or armament pylon, and may iuclude either a zero or a finite launch length depending upon the missile characteristics. 14-2.2.9 Loadlng The missile launcher should be designed to facilitate fast loading during ground operations. The loading process should require a minimum number of precise locating and positioning operations by the awmament mechanics.
14-2.2.4 Acceslbillty Maximum accessibility shall be provided to the launching mcchanismn, tubes, detente, firing contacts, and electrical connections to facilitate loading, unloading, circuit checking, diagnosis of malfunctions, clearing of stoppages, partial disassembly while in place, viewing of all dials and gage marks, accomplishmert of all adjustments with the appropriate tools, cleaning, and replacement. Minimum maintenance and turnaround times are a primary consideration. Other guidelines pertinent to this topic are contained in Chapters I1 and 13, AMCP 706W201.
14-2.2.10 Aerodynamic Effecto Effects of local airflow conditions upon thc initial missile flight path can be significant, and shall be considered during the launcher derign task. Immediately upon release from the launcher, the missile is exposed to aero6-,namic forces that tend to displace it from its intended flight path. Missile response is affected by launch velocity, guidance syztem operation during the launch phase, and control surface effectiveness at the !aunch speed.
14-2.2.5 Firing Circuit Testing Th-. designer should provide a self-cortained firing
14.4-.11 Ssipenslo mand Retentioa Suspension and retention omponents include the equipment used to attach the launcher to the air'4-7
706...2
X7o2o2
craft. MIL-A4591 contains a detailed method for calculation of suspension system iriterface loads. This specification is applicable to bombs and other
a dynamic and aeroo),mnic evaluation of the misuile/helicopter system. SouLv of *ection force that have been used succesdully imude compresed gas,
externally-mounted storm on fraed-wing aircraft. and
mechanical spLgp,
may be used for a helicopter misile launcher
devices. Provision sWi be made to prevent mad-
wlboe! rer the launcher is compatible with the lug and
vertent operation of the eWction system, eithcr in
way brace criteria contained tfrvin. Missile launch fixturm and suspension hardware generally should be 0 sAmple, lightweight, and small as possible, compatible with maximum reliability and with minimum ffects upon missile and aircraft performance. A safety lock or retention mechanism is required in the
flight or on the ground. W123 ROCKETS Rocket launchers of the number and type described by the governing hblicopter systeai specification " be installed The current rocket type
launcher to prevent inadvertent launch and to retain
qualified for use on helicopters is the 2.75-1. folding
the missile unde severe load conditions (i.e., cruh loads). The retention device should be designed to in-
and explosiv.
or propellant
fin aircraft rocket (FFAR). This roc~et is available in a variety of warbead/fuze combinations to suit
terupt the launch initiation system, as well as to rftrain the missile mechanically.
specific helicopter mission requirements. It is carr;;ed in and launched from the heAcopter by means of
14-2.±12 Lome& Iim'ilao
tubular launchen. The rodcet and some of its available launcher types are (kscribed in TB 9-1340-
The missile system should include a means of
201. ARMCOM should b:a consulted for (etails
transmitting a launch initiation signal from the nir-
regarding launch recoil foros, exhaust blast enve-
craft to the missile. The nature and complexity of this system will depend upon the type of missile to be
lope, firing power, and other pertinent items.
launched. Some missiles require only the ignition of a rocket mator, while others require in-flight prelaunch checkout, initial condition inputs to the guidance unit, and multistage launch sequencing. The *Jesign should be as simple as possible, consistent
14-23.1 Rocket Lamber Istaflatdm The primary function of the rocket launcher is to release the rocket safely front the helicopter without disturbing the roket from its intended f;ght path. The initial flight direction of unguided rockets direct-
with a reliable and safe launch.
ly influences delFv.,ry accuracy. There#fore, the launcher design should provide ;'or accurate alignment of the launcher boreline with the helicopter aiming under all actical conditions. reference The effectiveness, safety, deployment and maintain.
14-2.2.13 Restramlag Latch latch mh i s.metimes ArestraiingA lthmechanisn•sm etms is
launch. prior tolach missile just required to retain is_ estrinig therocet imilr t the Thisis This is similar to the rocket restraining latch discussed in par. ! 4-'.3.7. The latch is designed to retain the missile under normal maneuver loads, but to release at a predetermined load created by the riotor thrust. The mechanism may be designed as per- of the suspension andsuspnsin retettion system (rpar. 14-2.2.11I). The (ar.14-.2.1).The ad rteaionsystm by armaeasily releasable be restraining latch should and unloading grourd during men: mechanics contain should design latch The loading operations. provisions for adjusting the release load in order to compensate forcompnsae wear components. te mechanical mehancalcomonets. war innfrthe 4 ed to prevent postLocking devices should be proviO
ar1and em ents aity 14off pars. considerations and csro requirements ability req 2.2.1 through 14-2.2.7 are relevant to both rocket and missile installations. Additional interface design considerationstalated only to free-flinLt rockets are cono rain wi s thin tp ragraph. tained within this paragraph. heiicoptersTheconsist a fixed, forward-firing, launchers forrearwardFFAR currentof2.75-in. d lsr . Thein tuie vensistof venting, omyn-breech tube cluster. The individual tubes may ba reusable or replacehble, or thmcluster may be expendable. The launchers normally are inline of with flighttheunder specified flight conditions. parallel to the axis (boreline) launcher stalled ightallationsh cher laun of Othe Other types of launcher installations have been used successfully in fixed-wing aircraft, but have not been applied to helicopters to date. They are listed
14-2.2.14 Forced Ejectl.. Some missile systems may require a means of ejecting the missile from the launcher in such a manner as to provide separation of helicopter and missile prior to ignition of the boost motor. The necessity and mechanism for forced ejecton must be determined by
here because of possible applications to helicopters of the future, and include: 1. Open-tube pod, retrastable into the fuselage and extended for firing 2. Restricted-breech, rearward-venting, with constricted or deflected exhaust
loading variations in the missile-restraining force.
14-8
"
~*W 3. C3Lad~b.oukgh in wAMM 4. T~d= V &4VM li.. S MWu ddmomr1SA dam or 5vis q Sm." soibd linmm.
Sý 11111111PW MW1krali Minig Wh~M411111611 a"bumd budswe ma .guhe WHIMuIsM49POnbnusNe
OWASIUM .pineM A~aMs
i~u m.mamp~~
(Palr ffidw bm W&me ofa M k
i
TMs ur of dosed-bow&h hinmb* ta u qusi .m ariwid duqms ibm~wa. 1~ Sea of dos 2.75-i.. "AS me wie bo demes or 66"e from Eaedmw ii tols PMMW or fro iuesudm otdielaswd
omio to prw psins laudo WNW.M WO ibm biespa &MAido hiit' Wau.
tubs
amedmusis mmd &wGUdpmmi be&ammiuid VAd dos awimam hlsMplsU smms*ef land bows in md'in bl. n"ud mi bssiib. cosibum w" to ubiW Widow"i WpPlM" The iuammluei" m 141111111at. wi hout n (PA.I) C12111 loads111kM1t
TomM1111111111 LEON 144-M copmi of witabmsng lAmimw t~r& fining of then 2.*'S-ia IWAR haew been conawaselof Am fabric o a binmackisig or cm6s'eu Ezpumdab mowial. been tbs. huwve been camistugied of almmisam, Mimo,.mad maeurial. p~atic~mpeigsgAWu or coated popr-W A9* aaatwleh twha ave bee. Miiy smied anil (cud uSswAbk fortnes posvose mazy be used for rockt Winach tubes, ptovided dues amunt as take of
aMmd Flbe'41a.
--
--
A
--
L:2-
-... L
L-1-,-
lawJ.Red 3.p
TMugo"~ k"*
"23A3
JOV"a
J
" be
.it~
Grime Softy
The rosket im~deaatiom aWm inmhdm a Snpsmi anincba order to pow"t rmadisal aimt tY cutouW Glreg~ during grouad opuaftum. The firing ckm ca iss by 13 be PIw~awld by einding g exar ner the huieupierptyon la din AW ameesi l960 fa mf*psa * wea.o yiero anhifl
h an
dcrw
A nauAthy-opertea override swwm Aski be penstallation nqimns dadto ermt goun dbckout of the Mian circit. rocket *Antm aicwa xItgrunin pbfrec Lmd mch 1) 2.. on-M poss ctasct. Ors da. actbealrocatet fbyiasin Forward-firing launcher shoul be mounted with cost-un sj Grudn cotat fii acua roke the launch tubes at the optimum aagk, for highest il tenerg kds ahe oiring crcunilthy m IowWt SYV=i accuracy for tactics to be employed. The cor.t- n un~t responding aitcraft Pitch angle also should be contain provisions for preventing hinadywrtet rocket evaluated for One-half fuel load and maiu am motor ignition due to radio frequency awggy. munition load at the mean combat altitudje. PNovision for adjusimnirt of thv latuncher elevation may be 14-2.3.7 RestriningI~d ai t ni 1 . reouired so as to accoaunt for variations in the Rigaht lanhgtbeb mnsoanpppiae ohattitude of the helicopter between level and diving meaai nsiofh ~an apusritnge MautThbe' lancing. flight. Adequate stru.-tural clearance shal b-, proloade intoth lauto.r osriigmckhanism r whnThe nimatcl vided to prevent interference of the rocket (including roddtno aunced. is rocket gadi the reminen mandmust adjafins) with ehny part of the helicopter (including within the lauiicher durai'g the loading procedures. ngi lonh Arer rocketth nglmesrdfo 7ca stoge duri The latch will pirvent aft movement of the rocket, longiw o r ncket the fo s extp cengelt mosuef 7deg alf will restrain the rocket from forward movement isan xamle.and din. oftheexitpor ceterir~ with a force equivalent to that imposed by Wowag.u dinal crash load factors. If a blast.-peratcd detent ii 14-2.3.A Number of Rockets used, the mechanisin will prevent both fore and aft The applicable helicopter system spe~fifcation will movement of the nonlburuing~ rocket, sod will contain define the number of rockets to be carried and the sea means for manual rcleas of the individuri rocket quence in which they are fiked. Lateral spacing of latch~es in order to permit rapid loAding and removal suvcccsivuly-l1red rocket& should be selected to preof unfired rockets. veni the mutual interference effects of rocket bleuts, fin opeming, and jostling of adjacent rockets fired W42.3J8 Firing Cemuacts from closely spacod tubes. Rockets usually can be The e!4:trical contacts will be designed to provide a ripple, fired with relatively shiort firing intervals if low-resistance palb for the electrical airuat %lW mscceme irounds an' &irWfrom tubes with adequate 3_-9
40
-
Of bfeieuw ' loa* reakeL TOM maisred ppu @ inimA ha dspwdt* ~~ a uf mm @1w f99 -A I wkhei 5iP& of M~ phuwai of puea. TMu eagaces *A ha d~ued to paM.dinim hn Miml d htW of the Wommb 14%6ahad dthui km lb re" Sbt usMin mto 6= Me
f ~
Won =M wou i p**Wuu io ha be qqW Su 064 *8ot dod miA &dkw ~ - is th einombap WOWu4 fabrW@"s. biameuadom, dwd*oA. apWMe~Ms a ft Oun. @Wd mr~ef of 6@ PWMr.
-WW
$day wkik i* WI&
-
~
elsafe
~
bo4rANOB
-~ -bw
or*oapu sym anO be ha w-
Ckdn~, wi& both coan"t OnvuMMMU=6 Wow-
hoal aoomakk sad MaMi kirn of '3thah is a Pedann~iardqinsati, mequsam i td d ind vl MwTWhyo for mo** ods~oda and bogi ow qw wN ha dd..nd by the ban, haboopow ays. Wed&Ameio. At a mimumno ths op.UaU cowft mid AMW aft*S~wa offt swd of xf~thI to be fumt &4by. in rks ori rim*l ai a PREsc tk" imwvAq Tb. finmq 1 A Woal AO ha *sige fkw
qan dadaimupIan MG a48.M WAMMi s an mUAMWdau in the &a ph=s by 06 opagejm. While *vwy alur mm ha m*d to adiav Ohw smunu kved of ambty eoammum with annagmom system duip ma opuratma wadm al amedbnas, this ('bjUJve aso SUM be oamdood With W44 to tb attaianma of Maximinmm taatd fAM;*iwowa. whmr r*awoua mahy cwriwi 6npoas major
%,I.&.* KsArh F&hING If hwang"s .t bckway"MWaTItera Wh md Wo C"va lsWaab OpWAV cc to KMU*"a the rthAe in. aastalhor. the 4wja A&M0I w~khta. Wlag arn sk km&d at airspeed up to the mauamwu akvmfA yinWa.The covens als "d prvi6& pvotcdos towaa Wsit. duit, sad o~a awvrc wtawmatal wnd&uF~~w to which the aiguiaft may be m'ajecad. Mw. d.mgi of the tanag "lu owmn that the franto pferrfl t-m-a to radousi wiachbkAaf and 1OWti clOeuAM .apn 044PMOfes. airCrft S ture oir skin, and control surfaca. The hqaammu~ met a be p=Wto esOUX the OagM Intfike dut Sn fag nlatevah that hsvc been uased SOODIUD amsatig 1. P014yeste rusan with iner falB-Mawia sumS as chcpe odkophama abr, ju rate io, hic and eb~w girn. fitkthms 2. Plticifprognmted papws 3. =FamWp~mc q~W p~p
tical Wricimacy, an sapobawk kvle~olsakmybmayh to be assabliabd. This lr..rasaM ha basa irpow a torOUgh SOqiawi amfa abal &sadWm Mnati4 OfU the iaWaOWO M.Uura wampAr yMA t iwro SMd 6ad Ua~ur iauyfis (FMHiEA) humid be conucte (Mrth Pin 9YSSsn/v~hbdg Mnrf*ai daKi. This amnoals ommdau tbc '4cs of both wbayts Mmhw gd parsooml tr~ou durimag bow opabatomul um and =suftoaean Ekta from otiver nabibty. maatmiabi~iy, systeM qp~bcab Saeq hazards mv dsriuud in MIL-$TDfl2 S ClaM 1,4ie NClamd %US Mn;I a!, Cla" Ill, Critical; in lmIq aotobi.Tcfcl anyawyi should balirify f"aiha &K%"~m their refthanu effec uaPoz 6b.lcoter and PaOMasi. compen. provmonas and dkc hourd ckiwflatson. When Class I!U :-r Chuin IV hu&Ar are idenied, immediate actia:2 must bs iiken to elitmite Claws IV &adtoa mizuivizi, rbim Ill iteams, wmjbiteat ý.rilk de.ta okbje ivw Systatu %eefty rquirwvots an d* == fulkwin C&ý . ACP -M3.
-r--4
1424A SAFIT COSKAh)4 14-4A2 FIL* IueuWnpMn Tis paragraph speC&*k ly O'Wt the inwomec 0 Fimrinterruptio" devisan. ate qtar ona gue fixdndUAMovable WL-poaf aYvI.AW with oGur hqij -UIrv) for UKe faswin$ ra&1GL copta C1tMOO. 1. To protna* apgaint firioL at WM4i part of the airProper system safy diWgn wAais inweporate tht rTAt unvwte. r~tor Wanding gea, eatmial "ve.s ap~Aacbk dam~ and tw*A k~m ritd" w i~uaQi dec. ran*-Machiu vAI4IdSWpG yo must iv-id 2. To proeact apainsg fixmu into Cat path of Wo-h plaIW% rruuliaic or 0004. ia"pwitkwsaapt* bifdfinnr osau.ý sezb va othe guu.nsrxkets cc 'van &mgp. To tw d~divfw,m&-aachin rakw =1*0eSnc~ psds Fh*p mud be isisatud witld syw -*Wy wo pm.1.To iwtwaqA ls %apos frirv4 when the mar %idsa lo*gm rAid oaais&a 11r~ lrfOwAOI 5i. vs A run c~trmi srtm moht hase awd te ble spa9( do theuao symtm. Thema&.,a thol~hbga Oe b"is *auk%~l qpMldw "~Ibt 14-10
Jpniiuffalaed 4. To inemWe do wiemp firi wba my other Goadiuions un GOmidwei to be beau. dm1.t par-6w of to the mk*s Fri inlrawpsus VMS" ny Au
tbs level dfM gheamsn ps adl 410 ot keseam U.Nr. TurfMpsy hadiua beass. am O as
asI sysiam kw~ to delurmias pa pOWMd~
and IN
rubdoombu to a parnsiully bazarduis sibutsam. Fin
itmwnipt aS coatiroed by wassrmxy haesed on specific
WW Greu W&gi 118011 GM 680aunqim an deohwcýc tbly fuaW nt pcmsi the psobw control head Anc SM& M~M aia d vims Mno~k Wou wL In "i sqards Wo umfo. pokulam ar phav must be placi upo failure arneayes o(fi weasteurpt dei. - "I ur empiwýsrs "a be wokaluW aloq with po~t ulm U o ra Wl4ineesg dsvim for added -.
a
h lsdk
bSCOMMunI
AhSut
w,
as
eWm "Whan~d eWm* hamo ~s w~he kVAng selko, bet pmkqpd a=Ms to~ % of 65-90 4 a aove am cam dmms to dourmd Mc~ow bandiaWsf ess. Tin des. of tapows kWI iWbmhmu
a~mhisimiv A WM of 95 O~lb toVW* rH¶nL sapomere w~lm 10 dSa be tel-mad for a*~ 1S w. Desipa dwitke es be Ons Wonnauie 41 noms poblims is cider to addeie ss "" The U.Aiiq of nown to ahes kmb m eseaplithsd br
1. Do~w leimamim
The contou fooWr ASa poaoivr M@mitAuii amimes f Iiakkq U" it tav"l in thoem. arIasjhe usainf opmab Otherwism em WOM lMU It mMMY.a a ama buit iALO this tait am OrderW prevau kis at a wi% swre, for seumpls, aotsdgem burrelis AS u d, M ft~cam aia "M~ twrttu, oeoeWW %a Whd aioaa %Lz =rz. Tft~a cow=our wMu be Hsite w146 the pakwedso 4WAsSe UWW agwl gem diiput~a ma~npagpmd. U at & "et dastum foe "h pgghbited bmm. rhe dup of the wapporting str~uciamemus be such as to awiid wmsaf definoOr~a if the S" gglia Wk QWgeg mg axies rot&
a
sovad..a~ am omfis hr a 9w a
L41A.3 Coabso Falawus
)
Iskebrdhadwa u
a fuiwiee of~L bkt
eni~mc. Wed use
of am* at *0W"PM
*rOu thA ON Of X =80 405115 timmou sedleMbi £woftq pwo* 2- Dampi MW 3. Assawtion at ndhmS of doe S~emby se Of hwsis bdua, Or 641 mi0"Od ehwOLw
Wumpontsimmud 1A xasteamm go puweam d 'ý~
ak
ZL -:&
? W q eIAO k IoMhr ~onw 7hU app4Acahlm mlss prdtmmt aegi lW4 is MIL-A-UW. Ales. WL. STD-1472 states 'lii *qapmiu A& am pware Soess a01ss of tbs. asaim& ela'b15 We's prescribed by HEL-$TD-S-1-63, AFK 1*0-3. DUS*~ %mcflcatiam S-1-106 "Udlast 620AA. fijW vdG&.J~iy MKL-SW-.740, or MIL-A4S1W*, as appboAbh. 7W naipe" of Dois eepo the USafy Aa4 ONhIv~MM of 1424. Nee Unsiasa Gve~1ualegunbamlsOn ~or ~prownnel is& tscnbod in par. 13-2.2, ANICP 706-21'I %Mfpoms Ca zs*W !satrdow seitaa~oe. Fi "in of CWtan45do 'y OeO n 5A exSA
_W
Ca ca
AMCP 706-203.
1.D&Ain to the wapou 2.Wualma inalfmactioua, w~iibthc fti stin poe 164AA D*&l 91p* Ilibility of misflre 4r of exploding aaamugkition. whichbf~(-l~ L. e ovAtl mvlt in donsap V, the hullopwe or injm~y to fired roians, and links fiva. a f44i we~on w&eCi be Persomnnel considsrei from a usaety design sma~int- Dthns 3. Inability to re-fie the, weapc~n unti naftr a coei rma(e rtnt e a skervable "o-offpeid 1. Blown iqo a*s heicopter sitrucure with damiaBurst-limiting e~k=k gsorally am~ dassafnee a~s gn asit tsrnperat~uu-Iinuting or tiu..hm'ting. A tem2. !k deaiected into asnteswas, 1igbu, cxwrnsl parawure-Maiaa device normally as located on the Mo, c guni barrel. A tao).lm tiit deviceca be~locate 3. Be diverted by air 0c3flts into the rikin Of ord remotely in dhir wespon control circuitry. AIMCOM totors or wntin inlets, causin pomable catastrophic should be, consulted fof informatiev aboutL burntfatr limiting equifrawtav for specific pans amo for details 4. Cause sunl jaw. roaderina the weapo system aboutcurrent twaruntinawg devices. inoperable
)141-2AZ
Ced&ph Noin No~w resulting froma waepos fin"t can affwc benSru. praotIJw asoveamo and daacouufcnt red-sas
ares.smi !ý dlinje*A. OntOooci i~Lis 5 Bmse Cerain .ropont provie suftwacmt dnbrAS ##sake velocity to preclwde debris disposal probilm whum
Nw "MNOISis imuNSJ
00 hOuttda bkffl9#
FETUCU
, Eh
us*d "&WuW.y.kehi 114 GN A ft 66V V& *4a Vuiem wuk. a&musm iMla NWii. kum 14OOWO9S behimpar. cw.. Prookift of dis tuh
I S43
LS
bOWN"M" be
a
ae*y
ou'rwsbi les eashmt
-.--
morn
%4Lr~s
.hw~~fwa j
4lcO
e
W
~~
0Vb6b
StA
%md
&aaa of diVW-prhka bo i
WIN&
I us tufto iws diuihe a n
bm~e aw kewmM is de ofussz do awaraft two go*~ 1K WIN*&e oodM) by awm wkf met~ inpo.W
rswMeI~ puft uam
oWgpes8"O-ý
bK
~Amy
based on *aepwimass dfth awein* aid pamm boo thUat. Thus the MiAia df4tWita 'JOL a.inn*fad saimic he a mAppk* wish timaat awovE~~wuW4aA7 TesS, Evi . Gm Mai~ul= biaty kwetm aWndeauwirmam, ,ib ne mo The OP06iab Crieria %Wr tciic W ovtrm11 45'iM" to at~ftaS v~ftOM ft"hl toSo.kiw'i tbw &oalaa,* dileusma a iaipprpm~g1plop limit tina ane omixuakid in IMIL4TIDUOD. Tahe&A s ~ ~ ~ ~ ~ ~ i.2 ~ ~ foamrbA4u~*uan~o~~ ~ 'e of an aira~ft to WahMMif gumal GuWan cmNive toni; *Niam on peruoond if "i o d Mie tua2w the crew copaisuawt. AdeqMW& Vt~arai.ily r~uctuu Amdopone m pou a, v"W id a or"u to avoid "h haaw-d.
M um
G un ba rey wlt c n aqso uk u t s
Acusmukawno
as cahti mrvbck.w ndt*
~
t le veadu eqa atep~
astund or forced air tblimwsr) vaiiL~tatioa wo miran~da Coamusi~ou leel bdo dikaporvl pp.
o
omW
amd tvn* Of "t
O*SIxUSWU;4
"ims 8iwLaaet vlimacis. MWn
and 0)~
MW
h
~ i bo si
lsaapbwa
the beahaviot of the wapei coimpieti and that of the of paw above 75% of the Iowa- eap&oivc limit. thkt durig m the& paalls Vuarbdo cm rduopsed 4t whi buan cain &xte' issg~ crprto gox8Iiei'b asm-ss m st nore cefiwicfieai by propar tiabystew deep techuidewd M&Msdoms. Dcs~n of a put4 pa cmxaidi ad ~gedfinl hoer coditons~ ~ niqme than the use of armor. Ans optimum ~amtgn is chatacta~ized by atnior to.ivig Limited to the crew itabe stg tuy si bMyma ofth pwozirntiy UKe in msat. Roior downwusb t-jAzy be wida~itced as a means uetoahv virbitysdcux burial. concenirtion. *uplica.. of mat dimtmr&I The paiue systc-risiWi Dr designed ~" caeclczxad asi..,,~.pr. .Ia .ttr 1
iv,
*
t'j W,Usia OptatiOcal1 for "Rune time after cwiation cf
fum' ana ntrsfmro ~ d sala eot ro unhire to aVOid trappiDA &OaSs. V9111i03tion or ram Di~pfication. as an exampic. is the opt~imal wethod air purglu systamnt should be designcd'so that air infor protectioun of Right control~ sy;ci.Ascae takta ar kv,.at*J away front tk--Sun muzzle in order probleins are thfis need to separate to be effmuiv,. and SunSki along with purgin~g air. of1 toprestaei intakeci how f~ar to extend the redondancy. Duplicate GOAetýSs otato exany poisible. cakebemttaken ~~ai~~r'~ti are cf un g~m~st. ~ running side by side wou.ld not redkie overall vuiltrcm)alititng sources inv~sth om f pwos. u Fwinald nerability signaificantly. since a sinj)c. nit might iuplure thith acto. Moreover, if botlh acts k.miniate at A sg.must be deterirsaned by &a.tusifiringu monitored aCtUatcn, thart aCtUAtOr becoies the vulnerL dc~c.4n vqiprnt.Commoun by pope by pope qsrirr~nt.sole (~eeC.On element. Shielding and concentrution of ni4;rpvnerads joascral~y go together. TI'e numbers of he*,.) struc14-2.4.8 Twia ie(mster Power Swaveb KILrtl mcsnuocs Rad components and subsystems Ihit TO peeveast inadw-Sne11i V.peratien of a rut-ote C011can be regarded as not mission-critica! are very I trol tirrel whil;. it i6 beaing are 'ced, a mastv xower hrrii~od in a proptrly desigoed heficoptet. The praca'~svito~ that is ac.ja&abl itow doc turm si Nerics% itrca wxc of ,nczatratinmg critical components within a Ang b- provided. amall vvtuine, und then sisselinZ them washlek 14-12
I AMCP 7o6-202 critical rismn and structural menbar, reduces the stafagbal pirobability that any single projectile will hit them, but it also may increae the vulnerability of the helicopter to any projectile that does penetrate the citical cot . A 4aign that buries w.rienutrations of sensitive components uaide the beL*pler core. or even hides thern behind a single layer of structural protection, may affect maintenance and servic-bility. Analytical procedares for evaluai•g these trade-offs are well known, but !he importance that should be anigned to the various•values in the analysis is not so firmly established, Ballistic resistance is the construction of critical components such that they are massive enough to defeat the stated threat; for example, a control rod constructed out of steel armor material. As opposed to this. ballistic tolerance is the construction of a component such that the projectile passes through but the iten sill functions: for example, a multipivot point bellcrank of a nonshattering composite material.
Since vulnerability reduction by inherent design is discussed under the appropriate subsystems, par. 143 generally concentrates on the armoring method. 14-3.2 DEVELOPMENT OF VULNERABILITY REDUCTION SYSTEMS Designing for vulnerability reduction, irrespective of environment, involves an analytical procedure that begins with a study of effects of enemy weapons. The designer then determines how the existing or proposed helicopter system couples into, or responds when exposcd to, these effects. The uncertainties in the first group of data are minor since weapon effects are usufdly well documented; but in the area of system coupling, they are very great, especially for systems in the early sPages of design. A ballistic vulnerability reduction program shall be conducted with the development program to assess effects of design concepts, and provide analysis and guidance for controlling and reducing vulnerability by the most effective means. A vulnerability analysis of the complete aircraft is conducted as outlined. The definitions, criteria, and general methodology have been standardized by triservice agreement. 14-3.2.1 Vulerability Analysis The vulnerability analysis is presented in the form of vulnerable area as a function of striking velocity for each threat and each category of kill. The key threats to be considered in this study area are generally 7.62 mm API, and 23 mm HEI, (low and mid-
intecrity threats). The striking velocities to be considered are muzzle velocity, velocity at expected engagement range, and velocity at maximum effective range unless otherwise stated. The categories of il11 which are mutually exclusive are attrition, forcud landing, and mission abort. Each is defined: 1. Attrition. Damage to the helicopter which causes the helicopter to crash and become a complete ki after the terminal balistic damqg occurs. 2. Forced Landing. Damage to the helicopter whkh "aum the pilot to land (powered or unpowered) because he receives some indication of damage (a red light, low fuel level warning, difficulty in operating controls, loss o" power, etc.). The extent of damage may be such that very little repair would be required to fly the helicopter back to base; but, if the pilot continued to fly, the aircraft %ould be destroyed. The forced landing kill category includes a forced landing at any time after damage occurs (within specified design minion duration). 1. Mission Abort. Damage to the aircraft which
causes the aircraft to be unable to complete its defined mission. The following procedure and methodology are used in :hc vulnerability analysis. I. Target Technical Description. Detailed information on the construction and operation of all the systems, subsystems, and components form the major portion of the target technical description. The description includes a tabulation of all critical comn ponents, listing their title and function along with the failure mode; and the cause and the effect of a failure on the components, subsystem, and system (see Table 14-2). All potential damage mechanisms are considered as well as secondary damage effixts such as fires, explosions, and leaking fluids. Scik drawings with dimensions of the aircraft configuration with locations of its major systems and their components are used to determine the presented areas of these components for up to eighteen (18) attack directions (i.e., azimuth, elevation 0,0; 45,0; 90,0; 135,0; 180,0; 225,0; 270,0; 315,0; 0, -45; 45, -45; 90, -45; 35, -45; 180, -45; 225, -45; 270, -45; 315, -45; 0 +90; and 0, -90). Normally, the number of aspects used are restricted to the cardinal views (frorst, rear, left side, right side, and bottom); more views anre used where required for improved accuracy. If a computer is used, these drawings can be used to determiie the input data for tGe computer. The drawinp also should provide data concerning the shielding offered to individual target components by other portions of the vehicle. Additional detailed drawings may be required for critical components. Examples oi" citical 14-13
AMP 70&20 TABLE 14-2. VULNERABILITY DAMAGE CRITERIA DATA SUMMARY
CRmCAL COMPONENT
"EFFECTOF FAILURE ON COMPONENT SUBSYSTEM AND SYSTEM
FAILURE MODE & CAUSE
TITLE & FUNCTION
A
TABLE 14-3.
VULNERABILITY TABLE
PRESENTED *
SYSTEM
COMPONENT
AREA (Ap. ftz)
F3RWARD FLIGHT Singly Vulnerable Attrition
Forced
Landing
14-14 S
HOVER MODE
Multiply Vulnerable Attrition
Forced
Landing
Singly Vulnerable Attrition
Forced
Landing
Multiply Vulnerable Attrition
Forced
Landing
4
components may inclku
aircaft will be
iedleud m a inlt " inm
d. Transmisuon oil-cooler c. Hydraulic moddle and iwvwr
biliUm aPPiM for thu uomp fomtal or amm of vwor m vipm. is the dcIlsim for a ptirmh vi
redmioiao
of coaditimW kiA rA bt) isut or-
h. Control linkage and bell cranks
bIoy bt is bemd as IM ulfdel w thtd&whcity Of the threm aftar it has pKRfrW ie sbieding
i. Actuatort j. Crew k. Drive shafts,
riateria. A oopawas may w"be om s d iel ed for s0me or all of As premA am for am a men do -ni I ld. If thsk coow, the of the view coi
2, Kill Der'mition. Attrition, Forced Landing,
tim am• of d
ow.-
pon
fw ibs" •
m
wrait be
diq*. W%= wow is
Misson Abort.
r
3. Kill Criteria. For convenkiec of analysis, the helicoptr generaly is dividod into distinct system (i.e., crVw, propulsion, fue, flioht controls, rotor.
dw #01M of & to Prolact Sifmft comwiote s o aPas mwkiuq Whm tOw ermw a= a analyzed foraged, behied-plaft eflts will be womed for,
power train, etc.). Each system contains components that are vital to the succesul opcration of the sytm. Dbmape to a single component is said to be singly vuhzrsjabk. Components of a *et are said to be multiply vuluerable i a given kill catqeory if dam-aI Io than (N- I) nembers of the set does niot result a bhdicptr kilal., but d=z; to at le-ast,,nt WS of the et does result in a helicopter kill. Although the primary vulnerability analysis is conducted for the aircraft in forward fliMt, the list of vital components AMU show the level of kill for damage when the helicoptra is in a hover mode of flight (we Table 14-3). 4. Probability of a Kill Given a Hit (Px IR) tor Each Component. The designer assigns the Pi IH's to all the components and submits the list for approval to Ballistic Recarch Laboratories (BRL) at Abedoe Proving Ground, Maryland. Px If's are rrinental firinas on subioct item. Omlike -,sodon ex items,
-
of t
the woditional kill probabies for " Ww* W ie or a OFu vW fr all orM e d the vu or vie. This udugctim tof edkimd kilt pte-
f. Fuel cross-over valve S. Fuel call mumps
"'-
rt
a. Engine ompromm b. Fuel contt c. Engine oil-cooler
. Presented Area (A,). Once th¢ conditional kill probabilities for each component have bmn acertained, it is necessary to reconsider the target technical description to determine component presented arms. One of the least complex and moot frequently Used means of obtaining presented areas of neicoptUr components employs a planimeter to measure them from scale drawings made for each of the principal views to be considered. Several computer proorams, notably the MAGIC (used at BRL) Program and the SHOTGUN program, have bee developed for obtainiag presented arcea, as well as shielding for thb components. During this pfocess of obtaining presented are&!, it is neomeary to consider the space orientation of the individual components and evalu- ate the maskinS or shiekhin! provided to each compont by the trot of the helicoptr. !? general, any
sheng provided to a particular compvmt by the
I.e., nmber, wi~t vdob
y. aWd dkmciem ofm
aw &USdltawibe (For menew typesof~arm.u 'aw~in sielis Piropi esM reirea) Data 00 "-Jim" containwd in the clawin~d do&as mual this sbof reports CommOi BtL o.omu $t una a. kiI i.re x An.e V..-nblc ,,., probabilitiesIba, bow Wmbmud for tm w0UPat eas of and the psuleed neats of the b ctat thime components bave bun desterined, vuhlmerahis Ara of these oaponanto am gweated.. For a Sim thrtt strikin VeocitY, and view, the vuinvrNh ame of a comonent is obtaine from Eq. 14-1. A. - A, .PF, AV -
A
H
(14-1)
vulnerable area of component presenmo arm of ompof'i component coaditional kill probability
PK IH Vulnerable areas generally are c•assified one of two type4, singly vulnerable ate= or multiply vulnarable ares for singly and multiply vulnerable components. 'he singy vulnerable Wa A vr of the system is the sum of the singly vulnerable aros Avi of d conponnta of the system.
A
4V i- I
. A F, • (Pli H), (14-2) I I--
Only a limited amount of subtolaling may be acnomplished in a systum of multiply vulnerable components IT,for example, vulnerable areas were cowposed for a multiply vulnerable aet of components forming a propulsion system whkh contained two 0engn6, two engne traammoit s avul tWm drive 14-15
dml& Mlfku•ti6*lco Womi
m t
os•we ot
kuin l^ thd= tdo v Mdn wan for tds n. W40 nM tVWK aw mad & dni, dl& tpvb be a ade nd the vuleerale aim•s
w do kt wo- Vw wud be a tet Ik
m ad left drie dofk
"L.The .wmot j I1l1
sdobeta wau-d
Stractow No d vuWOMBB WmMiu b. OW (ewar JmMrMl) 9. IstrumaM POdI. Come" nMd iiWWMMM huurmi urad.to tO. Cm y/WJldi~ws/Dva
-
fewo hWepMAt
a.
*wt ammuha la tpahe mbe (LAL. nadal e d
• Pb
TrUm
'ue.
ril Ihel do&doems
lpO-
dai uomoedd fPems iom b. rM plese oAMmad wiafow For wmrs
It. Rotor 3d Raedemdaea Spers VidMWMM hdkN*Ga ab-kr H"2 "•Tle parabi dchdNktlfuiu may bseunfl fro 12. hmimin S• dm d 'Sat m .".m am inoMowhoeed. Dies mn i•Pmame imdbm ~ moasd hiLe,iroomthiWOBathvettion had-md ft~ O J4,U• Vdamddft Dat f"m Sbe.
,r
Total sarcraft vulmerab'ity for al1 tlhras
a)
gaia v
a. b. Sea HdWMwow C. T-oo 4L COchk unrproba~lties
2.Fed Cdn a. Selminbat b.
Iatmun m.,
.
a. W demh.c/exthindshf
f. P)•iaw ram 8- Meet I"* h. Mqltipl tauks 3. Fadrlim . Seuli-seaet v b. Sbutofffitto .. Dupli n/eepmration d. Routing behcl structure 4. Esgie and Powe Train: .5.
..... -s
Fa.
44uIAuu s~ um PUMP u
b. c.
Self-.lediag oil resrvoirs Engine and transmison operation without
Slubrkaton
Dump tolerant shdting Damage tolerant gearboxes and transmis,-ons f. Fire detection/extinguishing X. Shielded fuel control 5. Finish: a. Insignia or symbols that could be used as aim points absent or subdued. b. Matte/camouflaged paint sch me 6. Flight and Engine ControhL a. Redundant mechanical controls b. Redundant hydraulic systems c. Fire resistant hydraulic fluid d. boot actuators jam proof and redundant 7. Shatterproof high pressure containers d. e.
14-16
and proinction amitios. Alre.. Avant C u Dwvdqmw
The objectivs of th aor 4
d. FEM fmld
- d_
amer thai wit facibicooddered and pesmd ia taw evalisato of the ma ,fc obnsid by sub,.quent chan.g ian compornt ar, locatim, or kill
WUA.2
ftn at-b.
itim diiotiwm6 kilk. ok. -h dhold be
tk. a
nf th
an
vmkay
desipr is to miniainia w~sd t.e. tL.
o1wr, ades111ly inuJn. other evemntia cew. Mdon~ requvmennot such as vifftrnalaiw
mon and mergamn qrw. Threoe sevral alternate armor configuration conceps should be developed in accordanm with MILSTD-238 wad the effects of each upon nircmew eliscopter performance should be evaluated similar to umthodoloy shown in Fig. 14-1. The configurtion that beat s~tisfes the protection need by producing the smallst chan ges in system weight, crew vi-ion, and crew motion envelopes while sko providing the a
.
na~a
,.f
..
na•i.
nn •
h
na
an optimum integrated armor sysum. Surveys of experienced Army aviators should not be ovcmlooked a a possibility in achieving a balanced confir.ration. The armor configuration that is most widely accepted consists of an armored net bucket, individually worn cheat protector, and supplemental airframe mouraed plates. 14-3.2.5 Armor Material Selection Testing and evaluation of armor materials are genera-ly unnecessary for the selection of optimum armor with the lowest possible weight. Refer to Chapter 2 for guidance on selection. All airmor materials sha/l be qualified in accordance with USAAVSCOM specification 1560-MULTi-001, ProCar~X1 Speci.•¢caion for Li.gwigh Ai,cr•t Ar."Wm.
._peDs
HMAN INIPUT VARIA..ES
. PERFOFANCE e VISION
S
AN [jrw014MEvC OsDBISIONS OF
PA•1AGE
AIRCREMENB | 5TH & 96TH PIERCENTILES)-
1
BODY HINGE POINTS
AIRaFT
* LOCATION _PERF
_______
0
*
EXCURSION REFOIENCE POINTS
raw
ORMANCE
MATRIX
__________
___AIRcRI
_
"PRlOTECTION NEED' MATRIX
AIRCRAFT INPUT VARIASLES
D• ASIC: AIRICRAFT:__________ 13ASICMFdTS/L T TALIMi TAT IONS IIEQU I•
I
JAIRCRAFT CONFIGURATION
(RIEW STATION GEJME)TRY CREW MOTION ENVELOPE
STYPE * MISSION * PWRFOIWANCE COMPONENTS: -..
AIRCRAFT SHIB.DING COMPONENTS.._ * INSTALLATION o DIMENSIONS s BALLISTIC (-ARACTERISTICS
j
ICREW CREW STATION:
•n'*
•
STATION VOLUME o rINSTALLATION DIMENSIONS
IREFERNCEEPOINT1 L0SEAT
0
CONI).•OS
DISPLAYS • INSTALLATION * DIMENSIONS CONTROLS * INSTALLATION 0 DIRECTION
J
0 SHAPE 0 LIMITS
F IXTURES 0 STRUCTURE d WINDSCREEN * EXITS S SEATS • S HARNESS SYSTT--WJ
"Filu. 14-1. Study Imput Variable 14-17
ARMOR INSTALLATION DEFS'GN CONSIDERATIONS Th. be*i approodacs to the armor type protection of kskaptes from ground fire consist of: 1. Placement of armor in very clome proximity to doe cosuponeat for which prtcto isrqurd 2. Incorpolivioa of armor int th veil stu 1mw or into the component structuletit 3 Conskidnrtion of benefits gained (if any) from g) effects of hclicopter structure the shielding (makina sad/or oomponents. TeebaSIC aproaches have been divided ivto dic prhmncipal ssive prowetion concepts -integra armoparasitic armor, and indigwous armor. Interalarmr cnsiaa o aror ncororaed nto Ithegarralts rmorecemnsits of or inoprtdit hdralic houin~, wich re (e~g,ctutc~ fabricated fromu mor) or for existing loud-carrying Th useof he itegal amor strutur. met obiou concepa is in the floor of an aircraft having relatively an/opazea sttion ofvulerabe cu~m~alae IQUS
V ots ntgral armor caii be POO~i~tt5give,
expected to be heavy,
lmitd mallr istalatins, andhusis geeraly for the' inhmjnit hish"i weight af inte-. lb. e55 gralarmo i genrall is hatit mealli. Th cermic glarmpor iestand ot ieeralnonmetallic. mTherialsameinc wcm psuite for ad h omtcmaeil r o wlsutdfruse as load-carrying incinbers because of their lack of strengh or low ductility. Also, becase rmonteral isby dfintio a prmaent inta~lationa. materials having poor multiple-hit-protection capability are not suitable. Permanent or inte-1 pal armor should be made from materials that have the capability to withstand multiple hits, and that can be repaired in some manner without the necessity for removing structure or components u'ars~ti i asrcnearorin omsarmored helicopter rtructure. It provides only ballistic protection, and doos not function as a load-carrying mem bar. This concept pro,. ides more flexibility in th~e seetoAfmteil hnd teitga n indipmnus, armor concepts. There are, however, poteintial drawbacks, including the need for sufficient structural hard-points for adequate attachment of armor, the necessity for strengthening of backup struc. ture; interference with access doors and other items essetial to the proper use of the helicopter; and the establishment of realistic tolerances to allow armor panel interchangeability in the field between helicopton of the same type. When parasitic armor is innstalled on the exterior of the helicopter, crashworthinews is not affected particularly. However, if the armor paneb arm located in the interior of the helniopter. design of the bracketry and attachments is dictated by a crash-load criterion. Convertibility to
peace-time operation is an attractive feature of a parasitic armor installation. Indigenous armor refers to the benefits, if any, that may r"sult from shielding of the critical cornponcrit by aircraft equipment or structure that normally is located, or possibly way be relocated, between the projectile and the component which requires protecThe auggested technique is to take maximum advantage of the sUtruture and equipment or the helicopter in providing ballistic protection for the item. This involves accounting for the indigenous prottction that is inherent to the original helicopter configviatioti, as well as relocating equipment to improve the protoction. Consideration is best given to ballistic protection in the preliminary configuration design 5taV. Under th-, impact of a projectile, certain items o~f equipment and structure will create fiagnients and splinters that, while they are not lethal individually, will farmi a pattern covering a large area, thus greatly increasing the probability of a peisonnel hit. Other
items will arrest the projectile energy completely and
thereby. 1010% ballistic protection. The fact that the net effect of indig-nous materials may be either ,..~..
sitive or negative iwiicatas ih i'is UIBUlky
u
a method for employing indigeious armor. Consult BRL concerning data relative to the ballistic propaeil.Rfrt ete fvros~rcua aeil.Rfrt fvrosnrcua etc Volume 11, Ref. 5 for the behavior of fuel cells as msig 14-3.3.1 Aireime Tony Armor Service experience with seat mounted torso armor (cheat protection) has been unsatisfactory and r=iWuted in the development of the individually worn vest. Armored seats and restraints shall be eindt nefc ih hs et n r hma thfrwdpoecinUSAm NaikDvlmeat Center, Natick, MA, is responsible for this equipment and details may be obtained from there. ecablllt 14-33. lb. Armor of identical location s"l be interchangeable between helicopters of the same model. Ifarmor for more than one level of protection is provided, both should be interchangeable (seeMLlSO)
14-33.3 Re.wsbhlky E-.'cept for tht integral armor, each componeiit of the armor systern should be capable of being removeed by a maximum of t-vo men, using tools normally found in line masintenance areas. Parasitic armor shsould be removable in order to permit inspection, rqwir, and maintnansaw and to
•
__
I
,
...
•
0_"_0
AMCP MM620 provide access to masked components. Consideration should be given to the removal problems associated with combat operations. The design of armor should take maximum advantage of the modular concept with respect to size and contour. The maximum weight of a single piece of armor should not exceed 80 lb. Where the working area is restricted by existing structure and/or equipment, the weight should be reduced accordingly. Transparent armor should be readily replaceable and cleanable. Removal should be possible without disturbing the fire-control system. Each component of the armor system which requires three or more hours for removal should be considered as structural or integral. Such armor should be evaluated as pan of the helicopter structure with respect to load-transmission characteristics and fatigue life. 14-3.3.4 Flying Qalltis The removal of armor, partial or total, should not adversely affect the flying qualities of the helicopter. The use of ballast should be avoided. S14-3.3.5 l,-uebgllmz i Immobilization, as related to armor, is defined as rendering any moving part of a structure immovable following ballistic attack. The three most common types of immobilization are burring. keying, and deformation. Tho edges of pieces of armor may become cracked or tom as a result of projectile impacts. When two armor surface. are very close together their torn edges may come in contact with one another. The rsultant burring often causes immobilization, because the S. .. e,.,.vges prent ,,r rovnm..._ Kviq _rek occurs when a projectile or frmgment becomes
wedged between two surf..ces, or when a projectile
__.
penetrates one surfAce and partially penetrates the other surface so as to lock or "pin" the two movable parts together. Deformation results when projectile impacts swell the metal or push it out of shape, thus jamming moving surfaces together or otherwise preventing normal operations. Protection cgainst, or insensitiveness to, these types of immobilization always must be considered in aesning moving parts. In somecases, the use cf projectile deflector strips prevmts immobilization. A large clearance between moving parts also offers pro-
tection against immobilization. 3 Ar cashould 14-3.3A6 Armor Materil Attadtmst/istallafa I-.33h~l Moutnatg of Armor Plae A-mor plate sWall be mounted on strong, rigid
ANN&9
structural members or on energy attenuating nmmberu that are designed to reduce the peek impact loads th:.ieby reducing the structural requirements and hence the weight of local airframe stuct-ure. Doflection of the armor she/i be considered in determining space allowances. Wherever possible, ,wountings should be on the rear side of the armor plate (remote from the anticipated direction of attack). This arrangement, which is advantagous fr~m the strength viewpoint, also prevents damage to the mountings from gunfire. Refer to Figs. 19 and 20, Ref. 5, for armor attachment methods. The structure and brackets for armor support should exhibit sufficient strength to withstand the normal loads of flight, gust, blast, and landing, as well as crash.noads. The strength of hinges, locks, and fasteners on doors and/or removable inspection pancis should be maintained at the design level. The added weight of the armor should not strech springs, damage hinges, or warp doors or pan& to which armor is attached. The attachments should retain suffici-nt strength after the limit load imposed by balltstc impact to withstand flight loads without breaking loose or incurring zdditional deformation that would cause interference with any critical component. The bracketry may be designed to yield under limit load. When armor such as hard-faced steel must be oriented in a particular manner, attachments should be positioned to preclude improper installation. The sizing of fasteners is determined most conveniently by gunfire tests. When practicable, the armnor attachments should be desigred to take the impact loads in compruson or shear, but not in tension. Armor should be attached at three or more points having sufficient strength to support the armor and
to withstand normal operational loads in the event one attachment point is shot away. The standoff space provided by the armor attach. ments should be such that any deformation and/or deflection of bracketry and armor will not cause interference with the functioning of the armor-critical component. 14-3.3A h"tintie. Design The installation of armor shdil not preclude the rapid egress of the crew in emergency situations. Motalk armor should ha•v no discontinuities or rough edges that muight et up strcm raisers. All edge. be broksn or deburred !n order to prevent delayed cracking. Ceramik composite armor should not be used for primary airvraft structure.
AMCP 7Ofr202 Shock-susceptble components should not be mounted on armor panels. Afttr ballistic acceptanice testing, no therm.al processing is permissible without reverification of conformance with the ballistic acceptance limiui. Deviation from this requirement .11be subject to review and approval by the pro-. curing activity, Armor should not be attached by methcid- that tzanswjit the impact shock froml the armor to the critical component. In general. the followi.ng detailed installation requirenients should be met: 1. Use flat plates except where a simple curved or bent shape is advantageous in gaining angular protoction or weight savings. 2. All armor installations stkould provide space for a possible future increase of 50% in armor thickness. 3. Avoid the use of cutouts or holes in any portion of the plate for supporting or clearing miscellaneous apparatus. 4. Do not alHow any cutting or burring after the fiWa delivery from the armor manufactume's plant since this may locally degade the ballistic capability.
MMtile
Any steel armor in such a position that it could direlt secondary or bullet splash fragments into vital components shall be provided with flanges, spall shields, or splash stripe (peripheral fences) to deflect these particles, Spall is the fragmentation of the anrmor, either on the impact side or on the reverse, with or without complete penetration of the armor. The armor matcrials listed in Chapter 1 fe~ature adequate spall ftsstaflcc. with the exception of the ceramic comnposites. The remedial action is to overlay the exposed ctramic facing with a bonded layer of ballistic nykc.n cloth and, where required, to curl tht backing up beside the tile at panel edges. The splash or spall produced on ceramic composite armor tends to form a rather narrow cone centered about the projectile flight path. REFERENCES
14-3.3A.3 Pu~es SpahA and Spell Bulle splash is defined as particles of the projecformed from the impact against armor. With steel armor the splash tends to travel along the surface of the armor, much like the flow of a fluid. Bullet splash to the cyesr and nlddcncued.ag bodies of the crew, and, isdangerous sial7atce
1. I 560-M ULTI-0O1, Procurement Specifiation for Llghtwe~ghl AircrqJk Armor, USAAVSCOM, St. Louis, MO, 17 Oct 66. 2. MIL-STD-1288, Aircrew ProtecrionRequiremenis. Nonnuclear Weapons Threat, 29 Sep 1972. 3. American Cmqferene of Government Industrial k Hygienisis. Committee on Thresho& Limit Values, Cincinnati, OH-, Revised 1966. 4. ADS-Il, Survlvability/ Vuin-rability Program Requirements, e 92 USAAVSCOM St. Louis, MG,we
to equipment. The effective motho& of providing Protection against bullet splazh are to deflect it away, to trap it, or to turn it back along its original course
5. TR 7-41, SurvIvability Design Guide for US* Army Aircraf!, Vol I (U), Vol. II (C) USAARDL (now~ USAAMRDL), Eustis Di-ectorate, Ft. Eustis,
14-20
9 15-
CHAPTER 15 MAINTEANCE AND GROUND SUPPORT EQUIMENT (GSE) INTERFACE with related climatic and drnvironmental fac15-1UC CTIONlong1 NTROD baRO IcON tors. Helicopter and subsystem installations
lion, ad rquiemets;thatmus befolowe bythe bnerfoaloed bytedeigns andgeqinemertintrestatimsfator hell ter twe h geicoeer to ansur satisatryqie interfancs be ground support equipment (OSE). The design guidelime presented herein arm based upop Army circuIars, regulations, and policy manuals related to operation and maintanenace support of supplies and equipment. GSE interface design must consider standard items of GSE and the maintenance personnel capabilities available within the Army's system, and must be in consonance with the Army's maintenance support policy. Specific maintenance and OSE interface design provisions should use built-in-test equipment (BITE), where feasible, and must imaprove efficienzy in the use of OSE at each maintenance level. Wh.en incorporated. these interface designi considerations
-
should gloves heavy be designed to permit persoeine wearting and clothing to peform maintenance in cold climates. Instruction markidga should be legible and cdsota they are viewed easily. T'he desi* 1~ emtgon evcn ymitnnepr sonc withi a maximum of safety and a minimum of skill. The dwsin must assure that personnel with minimum training and overage mechanical abilities csn perform the required servicing, miaintenance. and rcpair of the helicopter and installed equipment. Sharp projections that may injuze personel during operation or servicing must be eliminated. Vital comnponents must be prote:te to prevent damagep during servicing. 15-2.1 SAFETY Safety is ai principal consideration in the dasign of ,,.
t.
(P
tr.
aw VUM-c~unmctls inalum te t,-nC-requled ut )turn-around servicing, maintenance, and repair; thus, -/maintainability of the helicopter w~ill be improved. As used in this chap~ter, the tern-i GSE includes all equipment nwevrd to service, inj*.%ct, test, adjust, calibtrat, fault isolate, meas~are, aasemble. dint semble, handle, transport, safeguard, store, repair overhaul, maintain, and operate the helicopter andit ~~~installed subsystems but excludes personnel Oquipment, office furniture and equipment, and common production tools end toolivg.
acedesgnshall conaform to system safety criteria principles, and techni jtse as definWx in MIL-STD882 and Chapter 3. AMCP 706-203. The objectives arc maximum safety consistent with military operational requirements; control of bazards to protect personnel and equipment; and identification. elimination, or control of hazards associated with each system, assembly, or subassemnbly. These coniie o h sdrtossol nldbtntb following: 1. Hu~man factors
15- DEIGNCONIDEATINS ND 15-2 CNSIERATONSAND DESGN R~EQUIREMENTS The requireulenti for OSE interfaces must be considered from the inception of the helicopter design. Reliability, maintainability, serviceability and selftest features must be designed into the helicopter and its Installed subsystetms to minimize the costs of maintenance facilities and manpower. Human bntaors considerations, safety, and accessibility must be included in the initial design. Similarly, standardization of equipment parts must be emphasized. beginning with design inception, Designs should c'enform to the human engineering andi criteria of MIL-STD-1472. *principles Equipment arrangements shouid minimize the /need for removing equipment when servicing is performed. Debign considerations should include analy\\sis of the operationO~ deployment requirements,
2.Levei or training required 61 416eiericulls mild maintenance 3. Ch~aracteristics of fuels and hydraulic fluids, and their hazard levels during storage, transportation, and handling 4. Contuinment. of electrical and radio frequency energy and appropriate waroings 5. Protoction of pressure vessels and associated piping 6. Classificationi of ha~ards resulting from essential use of explosives. Particular attention must be given to possible -ridafunctions that could create hazards, and approp~iste design features must be incorporated to eliminate or control these hazards. Genersl rules to be followed in 'sfety dcsign are: 1, Clearances must be provided to pei'mit the intcrconnection of fuel and/or oil rill lines, along with electrical connectior.s.
.
15-1
A
2. Appropriate valves must bc providod to permit
3tallation of engine(s), transmission(s), rotor(s), and
safe servicing in cam where pimsure is maintained, as in pneumatic and hydraulic systems. 3. Electrical connectors must be keyed to preelude incorrect interconnection, and the installation must prevent arcing and exposure of hot pins, 4. Ground receptacles must be provi&d to insure a common ground and to eliminate any electrical poteatial difference between the helicopter and GSE. 5. Guards and warnings must be providod in situations where explosive devices are installed or where high electurical potentials are present or for other potmntially sudden and serious hazards. 6. Interlock switches and devices must be provided as necessary. 7. Connections should be quick d'sconnect type. vibration proof, and not require safetywire.
propeller(s). To the extent practicoble, the helicopter design should permit complete preflight inspection without the use of special stands or ladders. Integral nonskid steps, handholds, and work platforms %houldbe incorporatod to facilitate maintenance. Similarly, utility syftems should have quick-accees provisiorsi so that they can be serviced without special GSE. preferably from wround 3evel. Arraniements that require special tools or removal of other equipment to accomplish an aviation unit level interchauge (removal and replaceni-cnt) should be avoided. If it is necessary to place o--e unit behind another, the unit requiring !as-fr•vuent access should Ze located to the rear. Except for pro, tection, or other valid reasons, equipment requiring periodic inspection, serice, or replaorment shoulo not be placed behind or uider structural members or other items that arc difficult to remove or that can be damaged readily. Equipment should be isolated from sources of fluids or dir*. The designer should provide built-in check points to simplify the connection of fault-isolating tnsi
15-2.2 ACCESSIBILITY The designer should emphasize ease of servicing, testing, removal, and replacement of all equipment. Inaccessible and complex structural arrangements must be ,voided. Moreover. the designer should consider sectioning #.he helicopter structure for ease of n ntl
II
,IW
.
64^11r.
field. Except where weight, structural integrity, or stiffnous are overriding considerations, assemblies subject to pc.riodic removal should be attached with quaick-i nnect fasteners of an approved type rather than with bolts or screws. Tool clearance must be provided for installing and removing lines, nuts, bolts, and other fasteners. Where the use of tools is restricted because of remote locstion, temperature, or other factors, fasteners such a self-locking plate nuts or anchor nuts sbould be used to allow single sided
service and maintenance personnel :o accomplish their tasks without working in awkward positicis. Maintenance at the intermediate suppout and depot levels will involve the use of automatic test equipment (ATE) %-henthis ir feasible and cost-effective. The helicopter system equipment and components must contain the necessary teqt points to interface with ATE. End items of equipment should be designed to have the test woints required to permit performance evaluation and diagnostic tests - oonsistent with the policy of returning the helicopter to
If lubrication is required, fittings must be located where accas is possible using standard Army OSE without the need for special a-lapters. If positional adjustments are needed, the design should permit unobstructed adjustment over the complete range of component movement. Accessibility for the 5th to 95th maintenance personnel should be provided for items subject to proflight inspecions and servicing. Similarly. emphasis should be placed upon accessibility of all coinponents subject to normal maintenance. Such items include fill and drain plugs, filter elements, valves, switches, and other field-replaceable assemnbies. Large, quick-opening acces doors and ample space should be provided for servicing of engine accemories and replacement of components. The designer also should provide for quick removal and in-
placement of modules - limited piece-part replacement, and repair of designated direct exchaage (DX) modules by replacement of external parts or use of authorized repair kits. In tur., modules should have sufficient test points to permit performance evaluation and diagnostic tests in accordatice with Army policies 3nd practices for depot rcpairs.
15-2
1&-23 STANDARDIZATION Ground supl on equipment interface designs should consider two levels of szondsrdization; I. Equipment lev•;. Standard GSE Items (MILHDBK-300 and DA Pamphlet 700-20). 2. Parts and Materials lavel, Military Standard (MS, AN, AND, etc.), Military Specification Qualified Products Lists (QPL). Such standards should be used in preference to
-
-requirements
0.,,
special comnmercial parts or designs seving the same Human engineering specialists should interpret the or basically the same function. To the extent that perdesign to define HFE problem area" and provide formance is not compromised. helicopter design solutions. should permit servicing and maintenance at all levels Populations for desitn shoidd be specified. During with equipment standlardized for Army use. Electrical the iterative process of design. existing critical mainconnectors, fill plugs, fittings, and other interfacing tentince tasks &&WHl be defined and. when possible eliitemns should permit connection of standard GSE minated. withoqut the use of special adapters. AFSC Personnel Subsystems 1-3 provide zpacific Where items require modification in order to pr:criteria for human engineering design requirements. form the required function, the designer should speciINPCOTTSADDA OTC fy tools and tooling that will comply with GovernSYSTEM mental specifications and documcrits rclating to materials, prooet-Ps, equivalent tolerar~ces, and size. The need for increased tactical mobility reqvires efThe designer shodd eliminate, where possible, the ficicut methods for inspection, test, diagr-osit, and lor new or specdal tools for use on new prognosis in the tactical operating environment. A helicopter systems. design objective, when requested by the procuring acTh:-designer should reflect decisions of Optimum tivity, is to provide an automatic inspection, test, and Repair LcvJ6 Analysis, basod on Army maindiagnostic system capable of diagnosing mnalfuncteriance/support policy in the mnaintenance and GSLý tions automatically, warning of impending inlterfaae design. Equipment should be designed for mechtanical failures, minimizing manual inspections, testing at the indicated levels, giving consideration to and permitting helicoptzr components to be changad the in'tiface with standardized Army test equipon the basis of condition rather than o! time. ThiAs ment. objective may be modified by state-of-tbe-art, misThe design of individual black boxecs should per. sion, and wi~ight/voluwe restrictions. mnit their serviceability on the helicopter to be deTh automatic ianfte,-tinn. tzxt. and diaenmatic Sterminincd by meanis of &4li'CoQritainciid. flu/nro-go test system should serve the following Jemntgts: 1. Laies and accessoric6 circuitry. The circuit loops of multi black-box subsystems should be contained in a single box. This will 2. Fuel subeystem eliminate the nzad for calibration adjustments at the 3. Oi subsystem helicpter level and will permit the use of simple, 4. Rotor group, including main and tail rotor light-weight, go/no-go circuits. transmissions, propellers 5. Flight wuntrol subsystem 1154A HUMAN ENGINEERING 6. Electrical subsystem Maintenance parrsonnl having a minnimum of 7. Easential avionics training arnd relatively low skill levels. anti some8. Hydraulic subsystem times working under adverse climatic and ceiviranmental conditions, are employed to provide 'he needed service, maintrnance, and rcpa-- of the heliJAppropriate seat on should be installed percopter and installed systems. Therefore, designs must manently in the helicpter and in ma&jor system cornbe based upon the proper allocation of rnaaequipment perforniance for system operation, mainpnns sapc~c esi ei syicue but arm not limited to, temiperatture, pressume vitenatice and control. Thus, designs should all-ocate funcion ptimzeperfrmace, t~ iniize biation, acceleration forces, liquid flow, electrical operationsl cIonstrintadmmize opfrmnerating idfl con~tinuity, Ow~onivm chatraacriatica, and vapor deopertioal onsrains, nd ini~t oeraingand tection. Sensor outputs should terminate in a quick. maintenance costs. Humnan Factors Enine. rliag disconnect outlet in the v~cinity of the ground power (HFE) design criteria principles and practices may Wbconcr eialThsnorut wlitrfe found in MIL-STD-1472 and MIL-H-46855. Goals with (3SF capable of procesing sensor data outputs to achimv include: atmtcly L Achieve satisfactory performanicc by operator, Mtmtcly control, and maintenance personnel. 2. Redluce skill requirements and training time. 15-.3 P'ROPULSION SUBSYSTEM IN7 3. Incrtaso the relisbility of personnel-cquipment TERFACES *Aram(cicastoara
7)
combinations.
4. Foster design standardization within and among systems.
1-. 1..
EEA EEA
A prc )lsion
system consists of the engine or 15-3
ummshf air ladmsetlomi submysn, exhaust uubsystam. Nhs aad kubqkwtoi sutbowtms starlag subsystem, caroe,mamisulon subsystem. and APU (if appiloabe) Intarias dedgn coulatlons otadi per. IS-2 shd apply. Specific design attend.. must ojmctie sad In in. be Sivae to mn- tl~jt dlude qui&-chaagse capablsities and accessibility for visulmintnanc, Ispetion ad srvicng.Addikdal propulsion system duaign rCiteria am evaelable InChapter 3and in Chapter 8, AMCP M~6-201. 154.
WNTFWANGEABILITY/QUICKCHANGE(
&Wbe onle Opl emonetsandaccmmes O~l3 DP~iU~tspreclude u1813L cmpountednth engndce.sr Desgn obetivs ihanfclud cO~fP~f~fl SOIWOY mi araiinlt tht f~htaeintra
nablt.
Egie
omonnt
and ac
points as required to acooniplish diagnostic testing compatible with the overall elicopter test and diag ncetic system (par. 15-2.5). 1S4.5 OIFL, FUEW. AND LUBRICATION o ytmdanvle ~d lo rvt ri-i age clear of the helicopter. In addition, tim valves win be self-locking to proclude accidental loss of fluid. m eieo ulsriigae. hl emtes and thorough cleaning. Arma shall be marked dearly "for fuel only" and desigated as to type or types of furel as appropriAte. Filler connection designs will the accidental connection of water hoes with those carrying other fluids. All fuel tanks must be capable of being drained completely and purged
lncrdaugsbllty.....-~ WUPDUII ~emovale foms onactiee ki n ~aliit at hftail Instal om omoise~ maitenance costh gebt cat.Intech1 inc mluiuinasmareotintena quirmmentsISM ouliediDML-2.0 are Helicopter operational availablty is sensitive to engine chang req~ulrmena Therefore engne aceinry odues ackor houl bedespsedfoi at tachment, to the engine to form acomplete ongine as, Naubly. The deig of the bdw co srmug flow fbi installstlon and removal of the engine assembly as a aitenncecoss, nit Toreucelifo~ylo sinle slquni~oeduelfe-ycew~altennccaes. me of removal and reinstallation of the engine &asambly from the propulsion systemn must be aprincipal design objective. Quick-change capability should be achieved with a minimum of special tools.
while on the ground.
154.3 CONNWFORS AND DISCONNEC~T POINTS
1
-
--
Permanently lubricated hardwake will be used wherever possible. Required greaen fittings should be des-gne in accordance with MIL-F-3541, MS 1500D,
150,ad50. 15-3.6 GROUNDING A static electricity discharge path to ground must ~ ~ * ''UUI
'
landing gear, an exiternal ground wire. An MS 33645 rectptacle for grounding the refueling nozzle should be installed not more than 42 in. norleassthan 12 in. fr ~ " n airram m radeito gruelnd sythem fy roviseillronmst for thearframge to arony ly rovi sioids mur bter madepio oaywiiS ro the ia thog kiso 7SATN
A self-contained engine-starting system isrequired for most Army helicopters. rronsttngwl Externai electricai power frgon trigwd be furnished through connectors and receptacles in accordance with MIL-C-7974. The turbine jetstarting receptacle, if installed, will comply with MS 25018 if asplit bus system isused. The standard connection for DC is MS 3506; for AC, 400 Hz 120/208 V 3-phase MS 9062 wnill be used. 1-2 GON ETR 38 RUD ETR avmber of connecor and coupling typoses lf be The designer shall provide a 12-in, opening for a *lmInd Selectona of electrical colmecto" Ii deW ground heater duct (a6-in, opening may be used on ambed in MIL-STD-1353. engines below 350 hp) in the accessory section, forT r PONTSward of the firewall and close tc high-temperature AD !53AB"M rSJ.MNINSK~flN AD TET PONTSdemanding units such as gearboxes and oil-system Samors fincorporated within the propmlsion syscomponents. A heater duct opening behind the firete will piovid for real-time readout of engine(s) wall also isdesirable to facilitate cold-weather mainsod subsystam(s) condition, performnce,ý and eftonance and servicing. Aiciency. Sensor AaU be provided at additional Hot air should not be directed onto electrical barA-- ftnaad, quick-discannec mechanical and ecual oupinp ab beusmas required. Automatic shutoff couplings in accordance with MIL.C-1413 Aa~be employed with fluid lines to me duce spL.W and contaminaticn during propulr onm~a, manlateancs. For ease of alignment and mu-ijimmn quick-diang. times. MS 28741 fleuible hose, or the eq~ulvalent, s~oult be imed for fluid lines,betwema the propuhio syste and the firewall. MIL-11-25579 applies to bases for b er-ampaare locations. The
15.4
(
AMCP 7O2O2 n-nes or other equip-itent that could be affected adversely by high temperatures. Openings should be marked "Ground Heater Duct". 1.3.9 ENGINE WASH The designer shall provide a•x~s for connecting a ground cart to the engine watr wash system. I"-4
-
TRANSMISSIONS AND DRIVES
General requirements for transmissions and drives aie presented in Chapter 4 and MIL-T-5955. Installation design should provide for ease ot replacemeat, repair, and servicing, and should avoid reqsiirements for special tools and fittings. Standard eyelet fittings thould be provided for attachment of slings for hoisting components such as gearboxes and transmissions. Easy-to-read indices, keys, or othem markers should be indclded on gears and shafts that require radial alignment during assembly and installation, Major components and assemblies should be interchangSabie in accordance with MIL-l-8500. Accessories and accessory drives shall be located for eas of removal and replao,.i-ment. The desian engineer should make maximum use of standard fittings, avoiding adapters to the maximum possible cxtent. Sensors compatible with the helicopter inspection, test, and diagnostic system shall be incorporated (see par. 15-2.5). Oil systems shall be provided with a rimple, visual method of verifying oil level from the ground. Sumps should be equipped with self-locking drain valves and an adequate oil drain path clear of the helicopter stuicture. The oil filler cap should have positive-lock f•eawr.s
•andl should be• in a dt-an -ndn
arm.4 Co,--.•
sideration should be given to installation of a chain or cable to prevent loss of filler caps following removal. Required grease fittings sAu/ be designed in accordance with MIL-F-3541.
15-5
ROTORS AND PROPELLERS
Quick-removal features should be provided on the cowling, spinners, and protective housings. Rotor and propeller controls should be designed with quickdisconnects to facilitate connection uf control-circuit testers. Each helicopter rotor blade shall be identified by serial number and marked in accordance with Army Technical Bulletin TB 746-93-2 and shall be interchangeable within the rotor system. Rotor blade tip markings or equivalent shall be provided in order to "provide for ground-checking of blade alignments.
The designri ushal provide for sling lift of individual rotor blades and the rotor hub. Blade trat .lng and balancing techniques should be simple so as to eliminate the need for a maintenance test fliaht after tracking and badancing. The propeller assembly design sall provide for the use of standard tools and tool types. Also, attachment points hdll be provided for use of the cormbination propeller wrench and lifting device for hoisting the complete assembly. Blades will be numbered erially and wiil be individually interchangeable. The design should pent 'blade removal, ling, hadling. and installation to be accomplished with the individual blade it: a horizontal plane.
15-6
FLIGHT CONTROLS
The flight cortrol subsystem as defined herein inludes the primary and secondary flight control systems (if applicable), the routing systems, the nonrotating systems, and the trim systems. ROTATING SNSTEMS The components of the rotating portion of the flight control system should be ca able. of beins replacod without disturbing the rigging. A plunm or tapped holc in a relatively rigid member, such as i structuval beam, should be provided so that control surface rigging gages can be attached by a bolt or screw. The location must be accessible readily ao that maintunance/support personnel can perform the necesary inspections and adjustments. Care must be taker. to provide the work space neesusry fr the adjustment of pivot shafts, push-pull tubes, and b, ll criuks. When hydraulic or electrical boost devices are used, sufficient access must be provided to allow 15-&1
for instiection, servicing. and maintenancc,
If autopilots and/or stability augmentation systrcms are used, self-testing provisions must be induded to the extent feasible. Where self-test is not ftasible, test connectors for isolating faults of a failed component should be provided so that test oquipmert can be. attached externally. 154.2 NONROTATING SYSr•T_ Par. 15-6.1 applies equally to nonrotating systems. When cables and pulleys are used in lieu of bell cranks and push-pull tubes, a means of checking and setting the cable tensions to specified values must be provided. In all cases, wear points, such as beli-crank bearings must have acess for insp"-zion and cotrectioni. 15-6.3 TRIM SYSTEMS Interfaces for cable and pulley, or bell crank and I-5
AA pub puI tube, adrjusnets mut be provided so that
IS.-I COMMUNICATION SYSfI1WIS Intallations for the Interommmi.,ilcation
GO
munication radios, :Lgna light, and flanre Iiaic"
tre
IF f
play can be maintaned within q5P-
19,7 ELECTRICAL SUSYSK•MS
sK.
om-
must permit ready removal and replacesmn or comDonents
zxer~al power reavptaces sl be provided fer theuletrial Wbla~stli. ThUSer eoiglm atlon,
When self.test provisions am included, teot results must be visible to maintenance personnel. Ad-
or tovIal--s wtadamrds forbCrthoepw• uelta. pr've u for boni iud t edin he d ep . tlun pow• unit will be induded in the design,
justm t cortrols must be readily aceible. For equlpami that does not include self-test muu h ei hudpemtctra eteup
A ainimum vuriety of eieudeel connectors should
fleatuires, the deseg should permit extrnal teat equip-
be added only afte it has bee ascertained that tho
provided in order to simplify handling and removal.
beused.Cn~n ector selection wcll follow nst o nmeat to be connected readily. Plug-in connectors be uied. ConnMetor iition will follow imnstructons should be used thioughout the installation in order to cint ped ent in corc Pronne isions must be avoid breaking circuits for teot and adjustment purnmml to prwvmt incorrect cIL connections. Sfor lub in and " Samto4po11. Pnd troviions lbrhicatig wl and . dssl Iganprors Break points should be provided doe zo the transandelectric motors will be icluded in te . Amitter and antenna equipment. The designer shouid oreilbility to lubriating points, as well as for the inconsider the use of directional couplers to simplify turchsaw of gneator and motor tnrubes will be maintenance. When the directional couplers are used. -provide. calibration indicators should be visible without The battery should be accessible easily with no removal of the units. lpecal (0. 8 required for its remuovul and inChassis slides, runners, and tilting mechanisms stailatlon. If nckcadmium batterie re aed. the may be used to facilitate accessibility of avionic dsige Ami insure that water cannot be added to equipment When used, these mechanisms should be the battery without first tamoving the battery from equipped with devices for positive locking in both onQQaVG %a a w~immu"1 aum. if~ aUrn 1 Was= wAD tended and retrnctUd positionr. Handles should be battery is fully chU3Sd.) "Thevoltag regilator sbould be aiccssibl to allow for in the requi 9 st equipmt and adjustim the rguator to ths "mcribed voltage level. All Reaes ad dcrcut b•eskers W be v and udly as e The inalds aidv ad d lightin installati should provideawsy acoas for Wetalationof beor tubs Wiing rum a mud be dwipd to alow ow a for maimtens of wiwsjg. miiar m c nections to Iduticol tpae of cr•ulpwent on a given helicopter xould be med to avoid errors in wiring during nstallatfoa, newoea, and maintenance. If there is a posiblhity of ccomsing a SE cable conrtto to Incorr"c able, the cable onnec'tor keying shod prsudis snb hteconnwatl. Comrmon boading poees or connectors dks d be Iwovided for acnectiosa between th helicoptr wiring and the USE. Uhe badi deesn €ouldeations for the eectrical subaystem are dsQAlbed in Chapta 7.
15-3 AVIONIC SUBSYSTEMS Dea consWiations and requirements for avionic subsystems an described in Chapter 8. Only essential VFR mh commualcation/navigation avionksae considored in this paragraph.
1*-
Withdrawal should be in the direction of available spa. Cable lengths should be adequate to permit slide operation. Adequate sp"c must bI available for the inttallation and removal of mounting units and ocrws. Screws and bolts requiring accoss for maintenance should not be buried kider cables. Installation and test requiremanis for intercommunication systems and communication radio sets are contained in the applicable commodity spedfications and data. I 5m NAVIGATION SYSTEMS The USE inteface conaderations and requiremeats applicable to the communication system also apply to the navigation systm. In addition, con. sidewatons must be given to the necessity for alignlng the direct• finder antennas and indicators during maintnance. Sim arly. the gyrosyn compess t'stnmi alisnment requireient? must be considered in the installation design. Specific installation and test 'equiriments for the navigation systams are contained In individual commodity spUflsktions and data.
15M9
HYDRAULIC AND PNEUMATIC SUBSYSTEMS The hydraulk and pneumatic subeytan desi requirm ts arme diacused in Chapter 9.
15-6 -
_ __
_ _ _ _
'
,). .. ,.,
".
all
. ...........
1".1 IHYDRAULIC SUBSYSTEM TIle hydraulic gubsysema requires checkout connectlons f'r interfacing with ground test stands t n accordance with MIL-A-5540 and MIL.C-25427. The system design should include a means for bleeding and replenishing hydraulic fluid, The hydraulic system relates more closely to GSE than any othter helicopter systera in that the hydraulic fluid is transposed between the two in ths accomplishment of the basic functions of filling, bleeding, ,,id filtering.
15-10.1 NAVIGATION INSTRUMENTS The pilot's compass installation must indude provisions for swinging the oompass in accordance with MIL-STD-765. Compass correction cards for each indicator in t'm helicopter will be provided. Radio navigation instruments are discussed in Chapter 8.
M.-9.2 PNEUMATIC SUBSYSTEM Pneumatic subsystem ducting and coupling must be adequate for ground test purposes. The ground coupling should be included and appropriately marked. Where ram air turbines are used, an interface should be provided for driving the turbine by means of a pneumatic power source during ground testing. MS 28889 high-pressure valves should be provided to interface the high-pressure pneumatic system5 with the GSE. The designer should consider using engine pnemmatic starters, it the heWkopter is so equipped, as the ource of compreed air for the actuation of airborne components. By tapping into the starter duct, the number of ground couplings can be minimized.
15-11 AIRFRAME STRUCTURE The structural design engneer must include desip provisions for jacking, mooring, lifting, owln&. and transporting the helicopter by all modes of tramqort to include helicopter external airlift. Lift points or sling positions for lifting of the Wom plete crash/battle-damard helicopter should bG provided. Lift points must be designed to nect the gloads specified in AR 70-9 and MIL-A-8421F. Sectioning should be employed to separate the body and empennage, the body and wings, or the body itself into constituent sections. Each section should include clearly identified points for lifting and load bearing. Hoisting eyelks must be provided if hojisting is the prescribed method of lifting. Floor obstructions of any kind ,.hould be avoided. If these are unavoidable, however, removable flooring - of sufficient strmigth to preclude doformation in normal use - must be provided to permit inspection, repair, and replacement of underlying structure or equipment not otherwis acsible. All transparent areas should consist of easily-
15-10
INSTRUMENTATION
SUBSYSTEMS The mounting of instruments and subsysten components sihal permit rapid and easy inspection. adjustment, removal, installation, and diagnostic testing. Sufficient clearance should be provided for re-. v.--. ,r , p.. ... ,....,i. modules. Items requiring more frequent adjustments or inspections should be more accessiblc than th, -se requiring less frequent servicing. The design considerations and requirements foi the helicopter instrumentation subsystem are described in Chapter 10.
15&1.1 FLIGHnI INSTRUMENTS Pressure-a•tuted flight instru••ents All be in-
7 '4
.
stalled in accordance with MIL-P-26292, and thu specified provisionu for calibraticn Adll 1iwincluded. An attlmeter correction card must be provided when the comb.nel atatic pressure systen error and altimneter instrument error exceeds k 15 3 ft. When servo-operated ins rumt nts arc used, rnecessary interf'ces for clectica, tct equipment should be included.
INLET UI YT INSTRUMENTATION The desin requirements for the lnterf.'a of the serial vehicle subsystem instrumentation with the helicopter and the GSE ame described In CbjAptw 10.
replaceable panels. Removal and replacement should be possible without the removsl of other equipment. Bearings and universal joini. must be equipped with the appropriate provisions for lubrication, inspection, and replacement. Cowlinp will be designed for ease of opening or removal. Fittings as required for matinj with towbars and towing vehicles iuust be provided. Lugs and rings suitable for coupling with hooks used on towlirws, and standard methods of towing, also should be provided. Jig points fer treasuring and leveling in accordance with MIL-M-6756 shall be inclueed in the airframe design. Jacking facilities conforming to MIL-STD-809 sAMI be included. The desig ,e. .Al assure that each landing gear can be jacked separately w-thout inter. rence betw•ocn the jack and the landinS gear sys15-7
APMP 706-)20 tem. The wheels must be removable without requiring the removal of struts or any part of the landing Near structure. Towing provisions will be in accordance with MILSTD-40. The airframe structure must include mooring provisions consisting of lup or rings for attachment of m-oring ropes. cables, or fines. Whare the fittings are reosed, they must bave suffient clearance space for easy extension from the recessed position. The word "moor" ,iyll appear on adjacent exterior surfaces. If detachable fittings are used, the design should provide for dieir storage in the baggage or "toolcompartment. If the landing gear is not used for restraint during runup of the engine(s), fittings to withstand twe maximun' Joad imposed during this ground operation shi be provided. When installed, these fittings will be marked with the statement: "Attach restraining harness here during ground runup or
A tiodon lug should be provided on rach main landing Sgar leg. These lug. should pinuit cornpression of the shock absorber to a position beyond the normal static deflection point when the hboicopter is moored. The landing pear design should minimize the requirement for special servicing tools, equipment, or fittings. Jack pads should be marked appropriatoly. If the landing gear is not to be used as a restraining member duiing ground runup of engine(s). it shall be marked conspicuously, in accordance with MIL-M-25047, with the statement: "Do not use for restraining helicopter during ground runup of engine(s)". Landing gear design considerations are. discussed in Chapter 12.
Design requirements for airframe structural design are discussed in Chapter II.
In general, all design provisions included for aircrew actiation of controls also arc necessary for ground scivicin5. Additionally, sufficient clearances are required for operation and maintenance. Inshould have all proternal cargo compartments -
15-!2
LANDING GEAR SUBSYSTEM S
."
units of the ianding par shuuld u; mAumav, for inspection, servicing, lubrication, and replacement. All air valves should permit easy servicing When retractable Sear is used, easily removable covers should be installed on all exposed equipment within the wells. Mud guards and scrapers should be cleanahk- easily without remcval, and readily rcmovable for tire servicing. The attachme it points for the landing Sear must be designed to permit easy installation and removal, Retractable systems should include locks that require no adjustment, and are acoessible for inspoction without requiring disasembly of the actuating members of the retracting system. Each Sear in a retractable system should have a manually installed, lightweight, quick-release, ground-safety lock or pin The design of this device should eliminate the possibility of incorrect installation. The pear should be designed so as not to retract or be damaged if manual unlocking is not accomplisbed before flight. A red warning streamer should be included with the lock pin to indicate when the lock is in place. AD
154 't"
15-13 CREW STATIONS
•
U113sioa1s
marPTIL
•1
_
conU iCuUouy Lu avuiu puimu,
damage during loading and unloading. Cargo cornmpartment doors should open easily, have a positive means of remaining open, and provide minimum interferenec with loading and unloading operations. Whee external cargo-carrying provisions are included, their design should facilitate loading and unloading opera:ions. Designs for internal and external cargo-handling capabilities should permit standard material-handling equipment to be used. The desigr requirements for helicopter crew stations, furnishings, and equipment are discussed in Chapttr 13. 15-14
ARMAMENT, ARMOR, AND PROTECTIVE SYSTEMS
interface design requirements for these missioncsentiml systems are containod in Chapter 14. GSE ".nsiderations arc included in the spocific commodity specifications and requirements, and are no! included herein.
__
AMP706202
CHAPTER 16
STANDARD PARTS 16-0 LIST OF SYMBOLS
AIC
length, in - basic dynamic capacity, lb - load rating for life L, and speed N, lb - life exponent, 3 for ball bearings and 10/3 for roller bearings - change in interior clearance, dimensionless
16-1 INTRODUCTION Standard parts for the purposes of this handbook, are defined as those items normally used as purchased with no change or modifications, and manufactured to meet industty, associa.ion, or Governmental specifications as to size, materials. mechanical properties, performance, etc. The standard parts discussed in this rhapter are
B C C, e K
-
nonreversin, load factor, dimensionless
fasteners, bearings, ekctrical fittinjs, pipe and tube
K,
-
reversing load factor, dimensionless
L
-
total bearing life, hr
fittings, control pulleys, push-pull controls, flexible shalls, cables, and wires. The applications for and 0t-, limitations applicable to each of these are diacustad. Further discussion of standard parts is found in
L LV0 _
]
-laminate
- B-10 life, hr - B-10 life, revolutions -
Li
-
L,•.
-
-
-
M,
-
M2
-
M3
-
calculated bearing life at load P, and speed N,, hr laminate width, in. life. hr ratio of shaft internal diameter to bearaig bore, dimensionless ratio of bearing bore to outer diameter of inner ring, dimersionless ratio of inner diameter of bearing outer
ring to bearing outer ring,
AMCP 706-100.
16-2
FASTENERS F
Fasteners are available in many ty;eu and however, futeners for use in the design suid consttuction of aerospace mechanical systems hall be selected in accordance with MIL-STD-3I15. Generally,
fasteners can be classified as either threaded or non-
N N,
dimensionless - ratio of bearing outer diamctcr to the housing outer diameter, dimensionless - rotational specd, rpm - ýpeed imposed on bearing for fraction of
threaded, and furthnr as either reusable or nonreusable. Threaded fasteners indude screws, bolts, and related hardware such as nuts and washers. Nonthreaded fasteners include rivets, pins, quick-relent fasteners, rctaining rings, clamps, and grommets.
N1
- rotational speed at which the capacity or
iO-LZ
M4
t;mp r
P
-
P, P,
-
SF T t I V X Y
equivalent radial load is determined, rpm equivalent radial load. lb
- equivalent steady ioad, lb load imposed on bearing for fraction of
RKK
ILU IA5NERS
16-2.2.1 Screws MIL-HDBK-5 c.ontains allowable design loads for
all structural screws. InstallatiQn of structural screws should be performed in accordance with Chapter 4,
time i, lb prorated load fcr speed N,, lb steady load, lb vibratory load, lb radial load, lb
AFSC DR 1-2. Screws should be torqued to the maximum practicable preload, compatible with the applicable torquing method. Screw threads for structural fasteners should conform to MIL-S-7742 or MIL.S-S879. MIL-S-"879
shape factor, dimensionless thrust load, lb laminate thickness, in. - fraction of time, dimensionlcss - rotation factor, dimensionless radial factor, dimensionless - thrust factor, dimensionless
threads ;hall be u-ed on all materials wi.h a minimum ultimatc tensile strength in excess of 150 ksi or with a minimum hardness greater than Rockwell C32 or equivadent. These threads also shall be used in applications where the operational temperatures will exceed 450°F; in applications that require the consideration of fatigue strength; for all bolts and screws
-
P, P, R
rnm
-
-
16-1
i
of 0.164-in. diameter and larger: for high-temperature internal threads in excess of 900(F; and for threaded holes (other than nuts). MIL-S-7742 thruads, both internal and external, may be used for fasteners smaller than 0.164 in. diameter, and for electrical connectors. Screws used on helicopters should2 be restricted to two types of screw heads: pan or countersunk. In countersunk applications. a head angle of 100 deg shuld be used wherever ponibk; otherwise, a head angle of 82 del should be used. Self-tapping screws should not be used in the primary structure. They may be employed, primarily in nonstructural applications, when the use of bolts or rivets is not practical. The installation and usage of tapping screw s comply with the requirements ef MS 337/49. Scrw normally are muable, and are replaced only wben either the recess in the head or the threads have been damaged.
Bolts normally are reusable, being replaced only when the head or the thread has been darmrged diring removal or replacement.
16-L.. PA" The installation or removtl of bolts normally is a two-han4ed operation. Therefore. it is slower and more diffaclt than is an assembly using screws. Ano-ther disadvantage is the many loose parts - such as nuts, washers, and cotter pins - rcquired in conjunction with this type ol installation. Special care must bt taken that parts are noi dropped, later to find their way into the engine inlet or tojam a moving ameably. In general, bolts should be no lorger than neceary. When tightened, the bolt should extend at least two threads beyond the nut. Hexagoraal head be thrads and left-hand boltm are preferred, Self-locking boltsshould may be aoiedwhen possible.
vantage over nonfixed in that assembly and repair become one-handed operations. Nonfixed nuts preferably should be o" the hexagonal-head type. Wing or knurled nuts, which require no tools, Adll be used only for low-tension, nonstructural app!ications; wing nuts are the easicr to install. Nonfixed nuts, unless safetied, must net be used where fallen nuts can damage equipment. Sclfwrenching nuts may be used in areas where insufficient space is provided for maintenance and use of tools. Self-sealing ruts are required for fastening equipment to fluid tanks in order to prevent leakage. Nuts are reusable urtil the threads, or, when apphiprovisions, arc damaged during locking cable, theand removal reinstallation.
used in tapped holes when one surface is inaccessible, or when there is a requirement that one surface be smooth, and when temperatures do not exceed 250"F. Bolts are more readily replaceable than studs. Allowable design bolt loads should comply with MIL-HDDK-d. Installation and preload toqtnz requirements are described in Chapter 4, AFSC DH
16-2.2,4 Washers In general, washers are used under nuts to prevent injury to surfaces upon tightening the fastener, and to reduce the stress on the joint by increasing the bearing am&. Spacer washers may be required in order to prevent loatding of bolt threads in bearing.
1-2. Specifications covering approved standard bolts ar includod in MIL-STD-1515.
The imtended use and the temperature limitation shouii be considered when choosing a washer.
Bolts•.dl be installed .n such a way as to minimize the possibility of ics of ihe bolt due to los of the out. In control eystems, and other applications (primaMily in dynamic system:a) where lows of a bolt could cause a catastropbic 'ailurt, self-retaining bolts &W be ured, or trio independent means of locking or iafetying Adl be required. Bolts shall be installed with heacds forward or uppermost, taking into consideration ease of m-intenance and replaceability.
Washers shadl be selected in accordance with MILSiD)-1515. Dissimilar metals should not be used toSether (e.g., steel washer with aluminum bolt) when normal methods of protection against corr.sion. such as primer, may be damaged during ile ass.-mbly of the joint. Lock washers can be used to prevent rotation of the bolt and nut in nonstructural applications, but are not preferred. Preload-indicating washers may be
16-2-2.3 Nuts Nuts can be subdivided' into such general categories as locking or nonleckang, and fixed or nonfixed. Nuts shall be selected in accordance with MIL-STD-1515. Self-locking nuts can be used independentiy or in conjunction with such devices as cotter pins, safety wiring, lock washers, lo•cking compound, or selflocking bolts, as a mear's of keeping the nut tight on the bolt. Self-locking nuts should meet the requiremerits of MIL-N-25027 and hall be subject to the design and usage limitations of MS 33588. Fixed nuts are affixed rigidly to the helicopter chassis by riveting, welding, clinching, or staking; and are used specifically in ca.et where the thinness of the metal prohibits tapping, or where limited space results in inaccessibility. Fixed nuts also have an ad-
16-2
•t2•~~~ JJ I
-
Vý
m
w--
4
*
AMCP 7106-202
)
sed to pgag bolt preloads, but must be replaced each time the bolt or nut is reinstalled.
0tion
*
can bx locked safely with castellated or uclf-ocking nists. Spring pins should conform to MS 16562, and they may be used within the design limitations givcn in MS 33547. The use of swagod, collar-headed straight pins should conform to MIS 23420, and that of flathead pins to MS 20392. The standards for poeitive-locking, quick.1VleasU pins are given in MS 17984 through 17990.
16-2.3 NONCHREADED FASTENERS 16-2.3.1 Rlvets Allowable design loads for rivets are given in MILHDBK-5. Solid rivetsof specific type= and materials arc covered by MS 20426. MS 20427, M-3 20470, MS 20613, and MS 20615. Additional rivet standards, applicable to solid, tubulsir, and blind rivrts (both struclural and nonstructural) are listtd in Chapter 4, 1-.. mc-e~s atmr mc-~s athr 1-3 AF2DH 1-2.AF_ý,Quick release fasteners arc classified as rotaryRivets are a penranent type of fastener: they must operated, lever-activated, slide action, or pushbe destroyed in the process of their removal, and button. often they cannot be replaced by another of the same Te ye ffseesaeMtvl ayadfs size. Therefore, rivets shall wtr be used in any applima hs ye ffatnr r 'luvl ayadfs touse, do not always require special tooling, and are where disassembly is expected to bc nectuary durigifeof th nomal te hicopcr.recommended for securing plug-in components. Rivets shall be used in applications where they ar rlalon wisth-e sallo are po~nns ari cove rsQp"k subjected primarily to shear. To prevent ripping, the rlaefsear loaekona olo ae fsees diameters of ilhe heads of countersunk rivets shall be Specific requirementz for low-strength, quicklrethnthe thickness of the thinnest of the pieces release panel fasteners ame given in MIL-F-5591. The ladrthcfastn.Rqieetfo onebkga; thcyfasen. equremets or cuntrbiningar~ disadvantages of these fasteners are that their holding contained 'n Chapter 4, AFSC DH 1-2. power is limited and many typos cannot be used In a cas where it is impossible or impracticable to reah f te t. jintto bck ucksold rvet, bind whe~re a inico"u star~act is rcR~urcd. T~hzir advantage mach ~ on~ obc~ oidrvtbid ~ ~ ~ i" i bsim that theyofa canohe and released easilywih ihu omltaunadnw1l cbe attached rivets may be used. Blind rivets may be structural or ihu ~u.l n un tfoncomls. maiu acis generally expansion Mechanical nonstructural. shaped spindle through the hollow center of the rivet, Standards for stj'nCturs! and nonstructural blind rivets are given in MIL-STD-15 15. In installation of rivets, the distance from the center of the rivet hole ta the edge of the sheet dotrn~ds largely upot. the atitn's analysis of the joint. Rivets shall be lo~atad so that the edge distance isnot less than 1.5 times the rivet shank diameter, or than 2.5 times the rivet shank diameter on a lap joint. The design allowable load data fof cou~ntersunk joints in MIL-HDDK-5 is based upon an edge distance of two diameters. If lesser edge distanc'is rAre used, the allowatble loads shall be subeantiatod by test data. The bead angle of countersunk rivets shall be 100 +ldisg.
Fgreater
W62.3.2 Phus In tie rods and on secondary controls that ame tot sub;-xted to continuous operation, clevis pins may be tused. In these usagns, the reversal of stresses aod 'he chances of loosening are slight. Clevis nins shall not
16-2.3A Turnbekle &Wd Tem~absi In helicopter applicatio'w. turnbackles are used primariy in controi cabit uz5ripsadons, afid their uSO should be in accordance with the system speciracatior. for the specific helicopter me "d. In most cases, the desired end fitting (except a threaded tend fitting) is swaged onto the cable. Swinging is discussed in par. 17-4. Turnbuckle terminals may be of either fork, eyt, or swaged configurati-an. Following installation and adjustment of cable tension, turnbuckles must be saifeid to prevent loss of tension. MIL-T.5685 describes one type of turnbuckle for aircraft application, while a positive-safetying type of turnbuckle is described by MIL-T-8178. General-purpose turnbuckle bodies (MS 27954) sAll be safety-wired in ac-
pins should be used in all permanent connectiins where the absence of play is easentiai. They
bodies (MS 21251) shill be locked in accordance with MS 33736.
~) be used wh~w tight joints are required. RDTaper
codic with NIS17731i(cutrsuak) fastners MS 177c ,po trudnghade MS (cuntedsnk y be73 and MSutua 17732,fastenrsd. ed.myb ue ssrcua nelftnes
cardance witft MS 33591. Clip-lock turnbuckle
EW-
164M. R@11010111 Rhgs Rutaining rings or snap ring s wected in accot4wn= with MI~L-STD-1515 -- may be used to retain be6,ýot nasu and. in limited applications, to retain pbia at boklt Tb. application of retaining rings to beanring and sesh is discussed further in par. 1637~ Beti ektarnul and iptarnus retining rings are "ailable as Isilmed or reduced cross section types. tye a a lower load Tb. radncsd craý os ka 4Wic capability and, therefore. is used only in locations w0hor the land is not critical and whore the retining rim is not required to maintain a tight installation. Wben a tapered cross section retaining ring isused to mainatin a tight joint, both the location and the wid~h of the snap ring groove are critical. 7hU externaW ring is installed in an approprittely located proovu in the shaft or pin. Internal retaining rings require that a groove be cut in the housing bore. The primary types and confignrations of external and internal retaining rings am' described in MIL5113-1515. Extmral retaining rings used as primary fastcenr saha'be safety-wired in accordance with MS 33540..
~
k.~A~hcuopter
mh*
tM18
els, or other structure through which they must run. Grommets are available in a variety of materials, misking it possible to select the correct material for vach application. Grommiets are wovered by MIl.-G3036,4-6491, -17594, ane6 -22329.
~~1
4
16-3 BEARINGS 16-3.1 CF.NERAL Design and selction of bearings for helicopter appiications demand the consideration of several factors; life requirements, loads and 9poeds imposed upon the bearing, available space envelope, and environmentl conditions. D~etermination of the allowable space envelope iclystef'tsepnthdsgnoabarg ss~.Mxmmadmnmmvle hudb obtained for the outer diameter and width, together with the preferred values within these malges. Toe nest, and most signlf-cart, dcaign considersLion is the required operating life of the bearing. e .. *an
16436 Clamp and Greaswitl Clamps should be used for holding wires, tubing. or boome that are to be removed frequently. Hinge clamps are preferred for mounting tubing or wiring on the face of a pane!. thus facilitating maintenance by suppoiting the weight of the tubing or wiring. For hlag, plug-in assetbliea, positive-locking clamp~s %%muId be used. Grommets hAll be used whe-ever necessary to prjotect cables, tubes, hose. aud wiring from chafing mmaraq~9 -
opened tlvisb a specific attion on the Paon of the user. is obtAined even before fte nutis attached to the bolt. Vibration, fot examnple, k buot sufficiet force to allow this type of fastener to fall free. Self-reaining fasteners of the impedaa - type arm coverod its MIL-343030, and selt-retaining. potitivelocking bolts ire described in MIL-B-23964.
1143.7 Suif-re~uling Fastener Self-retsining fasteners are used as a safety lock in area where a serious hazard would exist shic-uld a bolt be lost, or sl: ,uli a joint be broken following loss or failure of the threaded connection. Their use in holico"tr is applicable particularly to control systems and dynamic systems. Self-retaning fasteners are not to be confused with self-locking bolts. When the self-retaining fastentr is hnwteiW into the holes of two sterfaces to be joined, nechalical safetying preventi' it from being removed reaily. Thus, a semnipermainent joint, which can be
~
.
*L.
Pi
a
component service life, the time between overhauls (TBO), and the love', of reliability diat must be main taimed Because buraing failures, and, hence, bearing lives, follow a derinihe statistical distribution, beoring lives are calcullatad for a given survival rate. The bea~ring design life usuAlly is given in tern~s of B-10 lifc, which represenits the~ operating time that will be exceeded by 90% o( the bearings under given conditions of load and speed. In bholoptef applications, critical bearings are designed for R 3000-hr minimum life, with the aciual value depending upon the c sysem peincation requirements. In determining the required B-10 life, the designer must realize thut the equations ased for life calculations are based largely on data from fatigue tests of bearings run under controlled and nearly ideal conditions. Thes equations do not take into account such adverse operating conditions as severe thernial or contaminating enirironmeirts, lubricant deterioration, or shaft misalignment. Where the potential for such conditions exists, the use of a higher calculpted bearing life ix appropriate. The loads and speed imposed upon a bearing are major considerations in determining the type and size of bearing to be used. Purely radial load- may bc carnied by cylindrica roller bearings. Combined radlial and thrust loads require the use of ball bearings or tapered or spherical roller bearings, each of which may be used either alone or in combination with cylindrical roller bearings. For ipplicat ions in
(
)
uf
AMC" 70O-202 which only thrus lods are pramet. Vpecial thrust bearing geometry (Rfts. I and 2). When comparing hearings are available in both bill and roller types. cataog capacities q,' oted by differ-ent manufPcturers, For high-speed applications, ball bearings or cylinthe designe~r must make acrtain that each capacity is drical roller bearings should be given first considerradefined similarly in terms of life and upwd. Load tion sin= their speed capabilitiks are sigfirmva-ty ratings based upon lives and speeda other th~an 50) hir greater than thai. of tapered or sphericel roller and 33-1/3 rpm can be converted to basic dynamic bearings. Where the motion between two maubern a capacity C by: oscillatory rather than rotairy, consideration should I be given to sliding spherical or jout nal bearings or to C .. C, kiy ,lb (16-3) laminated clastcrtneric bearings.\ 667 Environmental conditions that must be considered where include oly.rmting temperature possibility of conC, - load rating for life L, and speed N1, lb mainination, and corrosive atmosphere. Knowledge or LI-life, hr the operating temperatures to be encountered will aid in the definition of the bearing materials to be used. N, - speed, rpm Bearing steels usually cam be stabilized thermally for Thr equivalent radial load P is definati as that a particular rangi of opeirsting temnperatures. Enradial toad that yields a bearing life equal vo the life vironments that can cause contamination or corroresulting from the combination of radial and thrust sion may require that special sealing devices be incorloads actually imposed upon the bearing. Itis porated into the bearing. Operation in highily corcalculated by the re~lation: rosivs. atmosphercs may require the use of corrosion-PX R Y b(64 b(64 P-XR+Y resistant bearing steels. %.mce the design requirements are defined and the where possible bearing configuraitions selected, the desither R - radial load, lb sbould perform the necessary ba~ring life calculaT - thrust load, lb tons. For rotating. rolling-element bearings, the A'. Y - radial and thrust factors, repectively,I. r. V - rotation factor, dimenr.ioqlas -
1666 N
((i \P)
hr
(1&~ 1)
orRcfs. L'10
0
V= P]
IcV rev
(1-)
where L,0 = 16- 10 life, hir L,= B-10 life, revolutions C =basic dynamic capacity, lb P -equivalent radial load, lb N - rotational speed, rpm e - life exponent, 3for ball bearings and 10/3
The radial and thrust factors X and Y Arm based upon internal bettring geometry and are given in If I and 2.The rotation factor Yis equal to I kO teinner ring of the bearing is rotating with raspect to ht radial load, and 1.2 if the outer ving isrotating with respect to the radial load. In most helicopter applications, tie loads ind speeds vary over a predictable spectrum. In such instances, the equivalent radial load P can be expressed as a prorated load for a given speed P,., and is detei mined by: [le -
[1t'
for roller bearings The basic dynamic capacity C is defineod as the radial load that a btaring will endure for a B-10 lift of one million revolutions of the inner ring, with the outer ring stationary. This capacity cmn be defined equivalently as the radial load that will yield a B-10 life of5W hrat a shaft %peedof 33-1/3 rpm. The basic cynamic capacities of their standard bearings are available from the bearing manufacturers. For new designs, the Anti-Friction Dearing Manufacturers Assocation (AFBMA) has established procedures for calculating the capacity based upon internal
(
1
lb (16-5)
',
1N,]
where -N
N, Pi N 1i,
prorated load for speed N, , lb speed at which the equivalent radial loaý is determiped, rpm load imposed on bearing for fraction of time 1j, lb speed imposed or.bearing for fraction of' time t,. rpm fraction cf time, dimensionless (the sum of fractions must be exactly D~
-rotational
16-5
--
--
ý-1
~
.-
~
-.--
An 4ternate method of calculation Isto dete.'tnine the bearing life for each condition of foad and speed by using Eq. 16-1, and then to calculate a total beating, life br. -.
hr(14 r(16
4
~the
where I. total bearing life, hr calculated boaring life a. load Pi and speed N1, hr Bearing life at high speed mauy be reduced s4nifltandly by the effect of centtiupi loads imponed by the rolling clenients agsi~at the outtr rAXe. Tl.. calculation of the internall bernfiatl irses and bearing life is a lengthy procedure that is covertd adequately in Ref. 3.However, for the vsat majority of applications, the centrifugal effects can be. nqleacted. and. thetv~ore, Eqs. 16.1 through 16-6 arm valid. The effect of centrilfvgal force, howevcr, must be taken hato aceouit in the came of hither rotational speeds. such ia thosa found in some engine reduction gearbox bearings. Usually, the life of a particular bearing can be inamurna by a signifum-Ai factor Pimply by using Waring materials that offpr greater uniformnity and better fatigue life 1taaa doe~s w adard steel. Table 16-1 shows approximate life adju.amcnt factors for seveal superior bearing steels. Conizult Ref. 13 for a camples. discussion of hwaring life. The calculated IWlO hi~e should be multiplied by theme factors in order to efrom determine the actual 0-10life.to be expwW L,
IT
In rtany helicopter bearing applications, the load imposed upon Use bearing is vibratory or is a camnbination of steady and vL-'rato?-y loa&. Such conditions of iotdlmg may b-6 converted to an equivalent steady load for the PU,'p06t of bearing lift
cak-ulations. For conditions in which the steady load is xreaster than the amplitude of the vibratory load, method Illustrated in Fig. 16-11is used. The ratio
of vibratory to mteady load is calculated, and the corrvspc,.idin1 rnonreverulng load factor K~from Fig. W6 I isntultiplieri by the *Aual steady load to yield the equivalent steady load. A similar procedure is used for the canc of a reversing load condition (amplitude of the vibratory load greater than the steady load). nTh. reversing toad factor K, from Fig. i6-2 then is mur-Aplied by the imposed vibratory load to yield the equivalet steady load used in the life calculation. Antiffiction bearings are manufactired in various tolcrance ranges or classe*4.precision. These class=s hi~vc been standardized by the Annular Bearing
r
1.4
-r
v
___-
(I T
Peq
JLI
.9Il
STAYLD/ STVEBADY LOAD LOAD
-EQUIVA.LENT4
g
used as the basisfo the AFUMA ratings, vacuumZ ~~depomd material now iscos~idered to be the stanl~~ ard stee throughout the bearing industry. TAbLE 16-1. UFI FACTORS FOR1. ANTIFRICTION REAP-ING MATERIALS
en
Ks Ps
MATERIALLIFE FACTOR TEALBEARINGS SAE 52100 A!R MELT STE EL SAE 52100 VACLUM-OEGASSED STEFL SAE 52100 CIV M-50CVN -
-
-
BLEARIG
I 3
1 2
4
3 -
I
.0_
0A 0.6 0.8 Pv/Ps, dimensionless Fgr 6I qevlu tnyLa o IM1-.EpaaStmyLdfo 0
-
K
0.2
Cembluutlem of Soeay own Vibratory Lead (Nou~resenlg)
1.0
TABLE 16-2. COST VS TOLERANCIE CLASS FOR ANTIFRIMrON BEARINGS G~RADE JCOST FACTOR
BEA41NG TYP~E ___
12-
BALL
1_
13 Ps - STEADY LOiAD Pv- VIBRATORY' LOAD Pq- EQUIVALENT LOAD
1
3 5
1.5
3
73.2
RLE -(CYLINDRICAL AND SP~KRICAL)
____
I
5S 41
(TAPERED)
_____ROLLER
21 0
~101
00
-
Peq
-Kv~v
0.8iC.5tion, 0.2 0.4
0
"P..
0.6
0.8
d
1.0
the mating surfk"s. Unless the intarferece fit betwwr- 'he maiting m besis greater thn zer undier au operating conditions, the bearing ring that rotates with respect to the load will amep. In addilighter prcs fits often ame secfe for the nonin order to preclude amoepng due to rotatlug iamtarin~g irto.I
-
5 Is
W IDIWUI
162. St~y ~adfordepedeat upon radial loWd. support stiffness, suarface Fl~r ~l~ake MW oprtn temperatute. ftef. 4 give a Figre Eqlvaeu 6.2 StadyLe fin~ish, nthad of calm~ladqn the required shaft fit that MWnCoinlmatea O ~4Od7 ad V~etOY ai theseulq acton. HoWVor, this analysis assumes Loads~dd that the shaft is solid, while amos helicopter powir EngieengComitee (lEC ani tic Roler train applications use hollow shafts. A goo4 up. Bearing Engineering Committee (RBEC), and are proximation of the required fit for hollow stee shafts de~ignated ABEC-1. -3. -5. -7, and -9, for ball as compared to solid shafts is given by: boarings; RDEC-1 and -5 for cylindrical and spherical refler belnijsand Claass4,2,3,O0.andO00for taptred Hollow Shafk Fit At X 1 (;6.7) Sldha Ft*d-less roller bearings. The n'zmericul symbols are listed in uI~~~~IpSli piii y Shafti Fitai I + K2
~diameter, width, radial and side runout, and also are
'p.
-
.Interference
*
indicative of overall bearing quaality. The use of the higher precision bearing gradcs represents a substantial increase in cost. For most heicopter applications, the. middle gradvo - designated ABEC-3 or -5, or Class 2 - should receive first consideration. Only in those case requiring high-speed operstion, extremely precise rhaft location. low nanout. or high reliability should ABEC-7 or -9. and Clan 3.,0.or 00grades be chosen. The increased costs resulting from the use of the higher-precition bearings an; presz-ted ain Table 16-2, fits arc used between the tmaonag anc the sWsf o~r housing in order to prevent creeping of th a bearing r ings anid subsequent frettint arW w*Lr~r
*
.-
~16-7
whm +
I -
2,I+M
d'es(68 (lu)
+ M22 Kj
-
d'kes (16-9)
ratio ofshIaftinternaldliameter to bearltg bome dlienesionlusa P42 -ratio of boarlugbore toouter diamuter of inner ring, dimensionless Sundard bearings are manulictured to provide a small internal ruinning cimearce when mounted with M,
the fits recomrmended in the manufacturer's catalog. Wises beavie,- rats (leas clearance, or greater interfmrnacs) an n-Aded to prevent creeping, a bearing with increased inteirnal clearance must be used in ord.' to avoid pruloadin3 the bearing radially. The change in inmernl clearance due to interference rats for both ball and ioller bearings can be calculated for astel Aialls or housings by the following equations: dk5
/~IC Shaft Fit UJAX 1 + K2)
tAlC Housing Ffit
M3(Kj - 1) K? 3 + K4
(16-10)
des(1-l
wheat Rqj, Kj:, MI. and Ml2 are as previouIsly defined, and tAlC - change in interior clearance. dimensionluas, and: I
3'radial 2 .
+
K3 -
dOess
(16-12)
3
I +M!
U 24
4
wba-eguided ratio of inner diamewe of bearing outer ring to beaiing outer diameter, dimensionless Ml ratio ofbecarins outer diameter to *he 4 housing outer diameter, dimensionless The changes in internal clearance due to shaft and housing fits are added. The lowest range of initial clearances *.L61t will give a positive running clearance M3~
th.-I
Additional information on the selection, design, and installation of both rolling elemenrt and sliding bearings can be found in Ref. 5. Dewings subjected to high loads or speeds must be provided with sufficient quantities of lubricant to mainta~n a satisfactory heat balance and avoid theymal damage to the contact surface4. Grease lubrication or splash-type oil lubriuation is satisractory for bearings operatirg it moderate speeds in areas where the heat rejection charactoristics of the bearing housing are favorable. For higher speeds - or in potetially high-temperature areas - jet, splash, or ariar oil lubfication usually is necessary, often with the incorpoation of an oil cooler. For heavy loads, the use of high-viscosity lubricants. or of lubriants formulated with extreme presusre additives, often is necessary. High-temperatunt applications, such as in engines or in high-speed 164
e~ngn redu.-tion gearboxes, may require the use of syqtheticobase lubricantxz of the eater type. Thene synthetic lubricants also provide the advantage of good low-temperature performance. The characteristics of many types of applicable lubricants, as well as descriptions of lubrication systems. are given in Ref. 6.
J
16-3.2 BALL BEARINGS T1he type of bearing 'float widely used in aircraft applicationis is the ball Learing. Ball bearings can be classified in three categories: tadial, angular contact,A and thrust. 16-3.2.1 Radial Ball Bearings The most common configuratioia isthe single-row, deep-rroove ball bearing. Ut iscapable of supporting both radial loads and light thrust loads in either direction while operating at high speeds. While most ball bearings are produced to ABEC-l toleiance&, higher pr mcision beorings, such as ABEC-3, -5, -7, or .-9, are available foi very high-speed applications in those areas where shaft location and rtn~nitt sire citical, The hearino- in th~e Uniher trades usually contatin better-quality retainers, typically of machined bronze. In addition, these retainers are by on~e of the bearing raca, rather than beir4g positioned by the balis., thus further improving the high-speed performanice of the bearings. Prelubricated radial ball ýexrings equipped with shields or seals are available for grease- lubricated applications. Because the radial ball bearing is assembled hy radially displecing the inner and outer rings and then packing the balls into the resulting annular space, the num-b*t*-l*Jr o~f
b!-.
in
I
I--in*.
;cH--^.
number of balls in tbe bearing, and hence the load capacity, can be increased by the use of a filling notch, a counterbored inner or outer ring, a circumferentially split inner or outer ring, or a fractured outcr ring. These features yield bearings having calculated capacities that are substantially greater than those of normal, deep-groove bearings of the same size. However, each of these tochniques has its limitations. A filling notch limits the ability of a bearing to support thrust loads, bectause the balls contact the notch. Counterboring a bearing ring results in a beating that can support thrust loads only in one di- ".ion; therefore, such bearings generally are used in pairs (see par. 16-3.2.2). Bearings with a circumnferentially split ring should be used only in applications having a thrust load sufficient to prevent the balls from riding on the split. Lastly, bearings with afractured outer ring should aot be used to sup-
..
_AMCP
port thrust load since this type of loading tends to spread the outer ring apart at the fracture. Double-row bearings are available for applications requiring higher radial-load capacity within a limited space envelope. Two rows of balls are held between singe-piece, double-grooved inner and outer rings, This configuration provides approximately a 50% increase in radial load capacity over the single-row bearing, and it available in both deep-gi oove and maximum-capacity types. For applictions where migalignment is a factor, self-aligning bearings are available in two types. In one type, the outer race is ground as a spherical surface, yielding a standard space envelope, but with reduced capacity. The other type employs a separate spherical outer ring While the capacity of this design is not reduced significantly, it has a larger outside diameter for a given basic size. 16-3.2.2 Angular Contact Bearings Angular contact bearings provide increased radial • : and thrust capacity, but limit the supportable thrust load to only one direction. In this type of bearing, one. oi shouider of the outer ring is removed ipmost Sompletely. The remaining small shouider serves to 2 copleely.Theremanin smll souaer srve to hold the bearing together, but cannot support thrust loads. The bearing is ass ntbled by heating the cuter ring and then installing the inner ring, balls, and retainer as a preassembled unit. This construction permits the use of a maximum ball complement and a one-piece, machined retainer, thus yielding both high capacity and good high-speed capability. Angular contact bearings can be made with a wide range of contact angles. As the contact angle increases, thrust load capacity increases and radial load capacity decreases. At high speeds, a bearing having a high contact angle will experience a large amount of ball spinning, with resultant heat generation. Therefore, for high-speed angular contact ball bearings, the
706202
u shown in Fig. 16-3(A), the beating set is known as a duplaxed pair mounted back-to-back with a preload equal to the gage load. Higher preloada are used to ensure that the nonthrust-carrying bearing does not become unloaded completely, which car. result in Dall skidding. Back-to-back duplex mountings are capable of carrying combined radial and thrust loads', reversing thrust loads, and moment loading. They provide a rigid mount for the shaft because the lines of contact intersect the bearing axis outside the bearing envelope. They also providle precise location of the shaft since all intern&: looseness is removed. If a pair of angular contact bearings is mounted with the r.onthrust faces of their outer rings together, as shown in Fig. 16-3(n), the bearing set is known as a duplexed pair mounted face-to-face. This type of mounting also can take combined radial and thrust loads. along with reversing thrust loads. It doen not provide the rigidity and moment-carrying ability of mounting, but will tolerate small back-to-back the amounts of misalignment. ali If a pair of angular contact bearings is mounted with the outer-rinE thrust face of one bearing against the nonthrust face of the outer ring u" the othr-c bearing, the bearing set is known as a tandem pair, as bearint i. C T his type o mottn is g .
(A) BACK-TO-BACK
contact angle should be kept as low as possible.
For appli.'ations in which thrust loads must be
carried in both directions, angular contact bearings ,;
(B) FACE-TO-FACE
often are mounted in duplexed pairs. When a pair of these bfarings is mounted with like faces together,
they become preloaded. Two bearings are preloaded "ifa0l of their internal looseness is removed when their inner &rnd outer rings are clamped together. Preload usually is built into a ball bearing by grinding the outer-ring thrust face flush with the inner ring while the bearing is loa4ed axially with a given gage load. The thrust face ot the outer ring is marked with an identifying symbol. When two such angular contact bearings afe "mounted with their outer-ring thrust fam togethei,
(C) TANDEM Figure 16-3. Mountag of Duplexed Ball Bearinp 16-9
• ".•. .._ , • .,
. ...
....
...
. . . ...
. .. ... . ............
. --
,-
-
•
..
.
,
.. ..... . ..
.. ..
.... ..... . ..
..
.
A!P! usd to carry havy thrust loads in one dirction, with the bearings shering the load equally. However, tuwd=m mounting doc not remove all of the imnatl loaenesa from the beanngs, and, therefore, pemits somse $bat float. 6.3.23 lus Ball Deariap Thrust bail bearings arn available for applicatioms in which pure thrust loads are to he supported at modewate speeds. These bearings afford a very high thram capacity, but provide no radial support for the shaft. Thrust ball bearings are quite linmited in speed capability due to spinning in the ball-to-race contacts, and they also are sensitive to misalignment. emuse even small amutunts of misalignment can Ssidt in h4,h internal contact stresses, a high degree of -'pendicularity must be maintained between the b"eirces and the axis of the shaft. Thrust ball bearings are made for applications requiring thrust-supporting capabilities in one or two "dlaw.tions. The various configurations available are shown in such catalogs of manufacturers as Ref. 7.
rollers running between two flanges on either the innet or the outer ring. In some casem, one or two additional shoulders are used in order to limit axial motion and to allow the bearing to support light thrust loads. Such configurations are known as locating types of bearings and are designated as one- or twodirectional, depending upon whether one or two sioulders are used. One-directional locating bearings have separable rings, and incorporate a single shoulder to prevent axial movement of the shaft in one direction. Two-directional locating bearings usc two race shoulderi. to provide shaft location and light thrust load capability ;n both axial directions. Locating roller bearings have design capabilities similar to those for non-locatin3 bearings, except for a slightly lower limiting speed due to sliding of the roller ends on the face of the shoulder. Cylindrical rolli r bearings afford the highest speed capabilities of -ll roller bearing types. in utilizing these capabilitiew, the designer must be aware of the special problems ussociated with high-speed operation. Operation at very high speeds, usually while carrying rather low radial loads, can result in roller I
.A
1*
twll
&V .
WI ...
""I
3
failure. In order to prvent skidding, several tech-
16-33 ROLLER BEARINGS
iques can be employed. The simplest and moat Roller bearings are used most frequently in appliinsure that a lasispeet radial load sufcations requiring higlh load capacity for a gliven spactemitireliable method is tooligcnac is prlwent. always contact rolling maintain to fiient include bearings roller types ofand various envelope. Becatuse this is not always possible, particularly when roller contapered spherical, needle, cylindrica!,The a havee rolling the only load imposed upon the shaft is torque, figurrtios. Nneedle rolspr bearinc 9irtint length-to-diameter rao o significantly greater specialized design features - such as reduced internal clearance, out-of-round outer races, or the incorporation of two or more preloaded hollow roller the that encountered in typical cylindrical roller rollng o ethelp maitai introd havben or concave use bearings roller Spherical bearings. have been introduced to help maintain rolling contact co'av rolling eleumet in order to permit operation withe rolaligngeements ieen odersfto herthop iong. and a constant retainer speed. However, bearings inhousing. uprigschearsuuayrqueexniv with misalignment between the shaft pand the tiuany require extensive dlpmrating such features Bath spherical and tapeied roller bearings can suptesting. and development limiting Their loads thrust of and those radial combined port With high-speed roller bearings, in which the cylindrical roller than are lower speeds sliding of end rollers n retainer is guided by a shoulder on the outer ring, the bearines aremloweruthanste bn becau of sliding between the rollers and radial growth of the retainer due to increased temperature must be considered. In order to avoid inthe guiding ribs. 1.3.1
Cyt•miJ Refer B
a
Cyindrical roller bearings are used typically in applications in which a purely radial load is to be upportWd. In moat cases, the rollers are crowned in order to prevent end loading and to compensate for mull amounts of misalignment. Cylindrical roller bearings are manufactured in a tandard grade designated RBEC-I for moat commereial applications. Precision-grade RBEC-5 beatinso are used for critical helicopter applications wher high-speed capabilities, and very precisc location and aligment of the shaft are essential. Cylindrical roller bearinp are constructed with the
16-10
~terference between the retainer and the outer ring, the initial retainer clearance must be large enough to insure a running clearance at the maximum operating temperature of the beain$. 16-3.3.2 Needle Bearings Needle bearings conprise a special class of cylindrical roller bearings in which the rolling elements are long in relation to their diameter. Such a design has the advantage of a very high load capacity for a given radial section. Needle bearings are manufactured in both full-complement and retainer types. The fullcomplement types provide high radial capacity since
J
,E,
the maximum number of rolling eAemn-ts is ued. Retainer types sacrifice some load capacity, but have the highest speed capabilities of any of the needle bearing types due to the roller guidance and spacing provided by the retainer. Because of the high lengthto-diameter ratio of the rollers, these bearings are susceptible to roller skewing and roller end loading, Therefore, they aft limited to lower speeds than are standard cylindrical roller bearings. Limiting speeds, as well as load capacities for the various types of needie bearings, art given in such manufacturr's catalogs as Ref. 8. Needle roller bearings arm produced in several configurations for different applications. The most cornmon type is the drawn-cup bearing, which consists of a drawn, case-hardened cup surrounding hardened and ground rollers. The cup acts as the bearing outer race and incorporates a lip at each end to provide roller retention. The shaft, when properly hardened and finished. may serve as the inner race, or a separate inner race may be provided. Drawn-cup bearings are manufactured in both full-complement and retainer l types.... ,,bh,., • .... "--..i... _. -_ nmn..le "',
row spherical bearings can cry high radial loads while supporting only very light thrust loads. Double-row spherical boarings also have high radial load capacities and can support thrust loads much higher than thove permitted by single-rov designs. In addition, double-row bearings can be made with asymmetric rollem, which reduce the roller skewin tendency. and thus permit the beaing to ofrate et higher speeds. Speed limitations and load capaities for typical spherical roller bearings we preeted in Ref. 9. In addition to the usual single- and double-row designs, spherical roller bearings are available in a special thrust bearing configuration. Ilis type of bearing has a high thrust-load capacity and can be manufactured with either symmetric or asymmetric rolling elements. Although spherical roller bearings do not possess the high-speed capabilities of cylindrical roler bearings or ball bearings, they afford a combination of high load and misalignment capacity that cannot be equaled by any other bearing type. Thes unique attributes are useful in many helicopter applications.
ment of hardened, crowned needle rollers and a steel retainer, also are available. The inner and outer races are designed separately, using dearances and finishes recommended by the besring manufacturer. The crowned rollers provide fairly even stress distribution along the roller length, and.iocrease the misalignment capability of the bearing. Needle thrust bearings use cylindrical rollers arranged with their axes positioned radially with respect to the axis of the beaing and held in a flat, machined retainer. The ro•lers run on flat races that are hardened and ground. Crc mut M- take," to locate the race surfaces perpendicular to the axis of rotation in order to prevent end loading of the rollers. Because the rolling elements of needle thrust bearnp are cylindrical rather than tapered, some sliding always occurs between the rollers and the races. Becauk of this condition, a generous lubrication filn should be provided in order to prevent excessive heat generation.
16-3.3.4 Tape,'d Retli to-In' V In helicopter applications in which both radial and thrust loads are high, as on a bevel year shaf tapuad roller bearings should receive first conaWdra6oa. Of all antifriction bearings, tapered roller barisngs offw the .best combination of radial and thrust apaciies. The raceways and rollers of a tapered rolhl bearing are conical in shape. If the lins of contact betwee the rollers and each race are extended, tey meet at a common point on the axis of the bearing. This geometry results in virtually pure rolling at all points of contact. Becausc t uheluls i,;pcd upon --- " roller by the inner and outer races (referred to as the cone and cup, respectively) are normal to the line of contact, the roller experiences a tnt thrust load toward its larger-diameter end. This loud is reacted by a shoulder on the inner ring (or cone) called the cone back tib. S'iding always is present betwon the roller and the cone back rib, resulting in the significant disadvantage of tapered roller bearings - their limited speed capability. At the preent state of technology, tapered roller bearings should not be operated with a relative velocity greater than 7000 fpm between the roller and the cone back rib. However, this figure is arbitrary and will be increamed as more-advanced taper bearing d•eigns are produced. Tapered roller bearings often are mounted in pals or in a double-row configuration with either a onepiece dual cua or a cone. Either the pair-mounted or
16.3.3.3 Spherical Roller Bearin Spherical roller bearings are used to accommodate shaft misalignment. They arc made with either concave or convex spherical rollers in both single- and double-row configurations. One raceway is machined and giound to conform to the roller. In the case of convex rollers, guide flanges are provided on this race. The other raceway. typically the outer, is ground with a continuous apherical surface. Single-
16-ii.
I
dou04ble-row design can be a4justed dwifin I*s~uftlaton to preload the beurlng axalaly. This pruload Is usd to insur positive axial location of -thesfrtand to preven roller skiddlW~I. helicopter power transinkselo applicittions, preload wanly Is a4uaeed by means cl a hardened sOee spu mounted betwee the two boaring cones. This specer Is gSround to a thicknes that results in a pre. -dolimrined rotational dying on the be:iu1 Waaiu~l*. 9est~s of beaning ananuracturing tolemums, the Anal gpnd~n% of the preload spewe isa ruklend ermo procedure and an oveteie pr sliould be prov"Ide with the assembly in order toievere tht t proper preoaed can be achieved.
MIL-543P and the applicable ia&.Aw* fr type M bearing wre uhma in Tab16e In addition, two group of rod ends rebrelow the Balanced Design Series we produced toesa abe rWeurmuents, of NASE6I. The applicable standerds pertaining to this spesifinaton arn NAS699 and NASGO. Spherical roller bearings also we used for alrkam. applications where eupselaly heavy loads must be carried. They provide good self-aligning capabilities and may be used for rotating shafts as well as in os-A dillatory applictions. Such bouaring11s re dwaubed in MIL-S-W4. The standards applicbl to that specifIcation arn preasented In Table 16-5.
16&34 AIRFRAME BEARINGS
16%U5 SL")10ING BEARINGS
..y
Seerael seie of ball and roller bearings are Various types of sliding burings are used In manufactured specifically for airframe control helicopter applications. The control systemu of applications. These bearings Ipnuses- high static-load sveal current aircraft employ spbeial buarlags of capacity. high tolerance to misalignment. and typicalboth the grease-lubricated and the sb-~actr ly include integral shields or seals suita"l for gr -em types. A typical spherical bearing isshown 'uFig 16lubricattion. These bearings are tailored for control 4. Spherical bearings also are my Ala"l in rod ends applications in which they must support heavy comn;4 binations of steady and vibratory loads under conBall bearings for appicaion amNDRD TifrmeLEro FoR &;aUNC.I airrae cntolappictios u~CONTOL ROD END BEARINGS of the full-vomiploinant type which provide high-loa1 capacity within a limited spaew envelope. Retainers l are not required since high speeds arm not a factor. 'TYPE STANDARD Thms airframe bearings are produced in both aitnular and rod and configurations in accordance with appicable Military Specifications. Annular bail M 15 OI HN bearings for use in control applications are covered by MIL-D-7949, and the applicable standards under MS 21151 EXTERNAL THREAD A the specification arm presented in Table 16-3. The -
MS 21152
requiramnets for ball bearing rod ends we defined in TABLE 16-3. STANDARDS FOR AIRFRAME CONTROL ANNULAR BALL BEARINGS STANARD TYPE(PRVIOULY)
/KP-B
(NS 2a'00) (MS M701) (WS 20202)
STANDARD
TYPE
MS 28912
ANNULAR, SINGLE ROW
DSP SERIES
N~S 2U206)
MdS27(44
DPP SERIES KSP SERIFS
(PAS 20207) (WS 26261)
NIS 27646
-
MS 28913
MS 27647
B500DD SER!ES DW, GDW SERIES
MS 28914
MS 27645 MS 27649 -
KP-BS SERIES AW-AK SERIES
MS 28915
16-12
_NENLTRA~
ROLLER AIRFRAME BEARINGS
MS 27643 MS 27645
L
hP SERIES KP-A SERIES SERIES
15
TABLE 16-5. STANDARDS FOR SPHERICAL
TYPE(PRVIOSLY
STANDARD__
IS27640 MS 27641 MIS 276A2
jM
HOLLOW SH4ANK
ANNULAR, DOUBLE ROW
fWIDE INNeR RACE
ANNULAR DOUBLE ROW,
-
TORQUE TUBE
for installation In control od asseblie. Journal bearings are used in ores, such as rotor head scissor aemblies, whom the loads and speeds do not require antifriction bearings. Sliding bearings a not subject to fatigue failures of the type that occur in rollingdeleent bearings, and thus they afford an increased
Gre-wlubricated siding bearings employ eithe stel-on-ateel members or sinted bronn mrunning on steel. both material ombintio are highly resistant to wear if they are lubricated property and frequently. In many helicopter applications (oscillating motions). however, daily lubrication of such bearingp is
materials, lubrication. loadWood. and environment.
as liners or insers on the sliding
dcgree of reliability. However, slidinlg bearinp exhibit wea characteristics that are dependent upon
STAKING
required in order to insure satisfactory operation. Several self-lubricating materials are used currently
r
ofates
bearings. The most commonly used liner mat'ias, are Teflon fabric and carbon-graphite.
Teflon fabric is relativdy easy to manufacture and can be bonded in both spherical and journal bearings. Teflon-lined bearinlgs are relatively low in coat and, if properly protected, can provide very good service However, the liner material is subject to deterioration
GROOVE BALL
as a result of exposure to water, dust, and oil environments. Teflon-lined bearings also are limited by the pressures and velocities tha: can be imposed upon
__I LINER MATERIAL
-O1UTER RING Figare 16-4.
them. Table 16-6 gives limiting values of pressure and velocity for Teflon fabric, as well as for other commonly used bearing materials. A PV factor (the product of pressure and velocity) also is given in the table; this factor frequently is used as a parameter for initial deign. it should bc emphasiz-d that the valuai given in Table 16-6 are approximate and should be used only as a guide. The use of any self-lubricating sliding bearing in a critical component requires that carefully controlled qualification testing be pe.formed. Unlike life calculations for antifriction bearings, those for Teflon-lined bearings, as given in most manufacturers' catalogs, can serve only as a rough approximation and must be modified by the results of qualification testing and service experience.
Spherlcal Aircraft OBering
TABLE 16-6.
MATERIAL SINTERED BRONZE TEFLON FABRIC CARBON -GRAPHITE: PLAIN "RESIN-IMPREGNATED
PROPERTIES OF SLIDING BEARING MATERIALS FOR AIRFRAME USri
"LIMIT SIATIC LIMIT SPEED LIMIT PV V, (PRODUCT OF STRESS PROJECTED AREA P, fp AND SPEED) psi x fpm psfpm
PRESSURE OVER
8500
1200
2r,000
60,000
201
5000-15,000
200-500 500-1000
200-500 500-1500
15,000 12,000 16-13
Carbon-graphite also has been used for selflubricating bnaringa. This material cornbines the lubricity and low friction of graphite with~ the good compressive-strength characteriontics of carbon. It is slightly more resistant to ad-ams onivronments than isTeflon and has demonstrated a wear life equivalent to that of a good Teflon rabtic. Its disadventages Incdude brittleness, which may be a factor if shock loads are present, and the relatively high cost of finished bearings, which is related to the poor machinability of the material. The Military Specifications and Standards that define the requirements for sliding belApig for sitframe use are preseted in Table 16-7. LAMINATED ELASTOMERIC BEARINGSmotion BEARNGSIn The rotor heads in most helicopters employ ball, tapered roller, and needle bearings operating with oscillatory motion. Such bearings represent a substantial proportion of the rotor system weight, and IW6-.
reuire lubrication and sealing. Laminated elsatomer
bearings, on the other hand, require no lubrication and are lighter than conventional antifniction bearings for a given load capaciaiy. Thcy also permli(tic type of oscillatory motion and loading present in most rotor applications. Theme bearings consist of thin, metal laminates alternated with thin sheets of natural rubber. The rubber is bonded to the metal by a proc..s similar to that employed in the manufacture of lip meals. Four basic types of laminated elastomeric bearings ame made. Radial bearings are able to support a radial load only, and permit oscillation about one axis, Similarly. thrust bearings support only thrust load,
and also permit motion about one axis. Conical elastomirric bearings are analogous to tapered roller bearings; they are capable of supporting both radial and thrust loads, and permit mingle-axis oscillation. Spherical elastomeric bearings are the most suitable for rotor systems. Thene be~.rlnis employ spherically shaped laminates, are able to support combined loads, and permit oscillatory motion In any plane. The lamninated construction greatly increases the stiffness of the elastomesic structure in the direction normal to the laminations, while maintaining virtually the same deflection characteristics in the plane of the laminates as would be found in a solid block of rubber. Tl~i rermitit laminated clastomeric bearings to support high loads while permitting the degree of necessary in rotor system components. the design of atlaminated bearing, the most important parameter is known as the shape factor. This factor determines the.lod and deflection capabilities of the laminated structure, and is defined as the loaded area divided by the force-free area. For a
single rictangular rubber laminate, the shape factor SF is given by: 'c~ 2SF L, +
dB)u
1-4 (6-4
where L.laminate width, in. B - laminate length, in. t - laminatz thickness, in. The shape factor for a laminated elastomeric bearing is calculsatd as the shape factor for a single rubber laminate multiplied by the number of
TABLE 16-7. SPECIFICATIONS AND STANDARDS FOR SELF-LUBRICATING SLIDE BEARINGS
SPECIFICATION MIL-8-8942 MIL-B-8942 MIL-B -8942 MIL-8-8942 MIL-8-8043 MIL-B3-8943 MIL-B-8948 MIL-B-8948 16-14
fSTANDARD IMS 21230 MS 21231 MS 21232 MS 21233 MIS 21240 MS21241 MS 21242 MS 21243
SPECIFICATION
STANDARD
TYPE
MIL-8--81820 MIL-8-81820 MIL-B-81820 MIL-B-81820 MIL-8-81934 MlL-D-81934 MIL-B--81935 MIL-B-81935
MS 141 W MS 14102 MS 14101 MS 14104 MS 21240 MS 21241 MS 21242 MS 21243
WIDE AMINULAR, GROOVED WIDE A1NULAR, NONGROOVED NARROW ANNULAR, GROOVED NARROW ANNULAR, NONGROOVED SLEEVE (JOURNAL), PLAIN SLEEVE (JOURNAL), FLANGED ROD END - MALE ROD ENJD - FEMALE
AMCP ?W202 laminates used. For helicopter applications, shape factors of 30 to 40 have been found to yield the bat results. Additional information on the design of this type of bearing can be found in Re(. 10.
W1%
16-3.7 BEARING SEALS AND RETAINERS 16-3.7.1 Seals Seve al types of sealing dcvices arc employed in systems. The most helicopter power trains and rotor ip seal. In this cmcommonly used type is the radiat figuration, a V-shaped scaling lip, which contacts the rotating or oscillating shaft, is held within or bonded "to an outer case, which in turn is pressed into the component housing. The sealing lip typically is fabricated from either natural or synthetic rubber, For oil-lubricated applications, the saling lip is "forced against the shaft by means of .a circumfe,-ential gerter spring. In grease-lubricated components, this spring usually is omitted. Radial lip seals are fairly inexpensive to manufacture, and provide reliable operation at surface speeds 3M, fpm. For higher speeds, a to approximately up hv••ur,•l-wlyarm lin toot ahnidel nl"•v-Av
Grtt ensiders.
tion. In this type of seal, the interface pressure between the shaft and the seal is low, and a hydrodynamic oil film is maintained at the interface, Helical grooves, indentations, or ribs are molded into the air ride of the sealing lip and are effective in directing the oil flow back into the area to be sealed. Hydrodynamic lip seals can be operated at surfac speeds of up to 10.000 fpm because of the reduced contact force and the maintenance of an oil film that lubricates the sealing lip. For all lip seals, the compatibility of the seal material wit" %1%,, 'i"W to M , ,t beta- , -""account.Many of the current lubricants, particularly those of the synthetic ester type, have a detrimentd effect on certain seal materials. Materials such us nitrilc rubber and fluoroelastomers are available for usc with this class of lubricants. Their use results in increased seal cost, but often is necessary in order to prevent deterioration of the seal lip and resultant leakage. For applications involving high operating teniperatures or shaft speeds, carbon-face sals generally are used. These seals consist of a carbon sealing face, called a nosepiece, bearing against hardened steel mating rings, which must be extremely fiat. Such seals have been used successfully at speeds and temperatures typical of turbine engine operation, Because of the very precise tolerance required in the Smanufacture of all parts of carbon-face seat, their t is many times that of a similarly sized lip eal.
a
Therefore, face seals should be restticted to applications where surface speed or tempsratures prohibit operation with a lip sea. Clearance seals - such as labyrinth and ring =Lb also are used for high-speed operation. They provide satisfactory scaling when the shaft is rotating, but permit leakage under static coonditions. Therefore, these sals should be used only in applications ;n which a static head of oil is not prese•t in the seal area when the componeat is not in operation.
For applications in which small oscillatory motions are present, such as in rotor head hings, diaphragm seals have been used succemfsully. Thee seals employ an elastomeric membrane that samns the gap between the o, illating and stationary mewbranes of the seal. The membrane undergoes torsional deflection under the oscillatory motion. As ,-lArr long as the elastomer and the metal-toi bonds rr,.nain intact, the diaphragm so! is able to operate without leakage. Additional information on various available scaling devices can be found in Ref. II. 10.3.7.2 Bearing R Several m.-thods are used to retain benring rings on shafts or in housings. They include lock nuts, snap rings, retention clips and staking. A lock nut is used to retain the inner ring of a bearing on a shaft. Together with the diametral interferme fit, the lo& nut prevents the bearing from creeping under toed4 and also provides axial location for the bearing. The lock nut may be secured with a tab washer. Tabs inserted into slots on the nut lock the washer to the nut, while serrations on the bore of the washer fit Into -,.t:an: o. t.1-th..ad -o.f-..h... thy preventing the assembly from turning. For lightly loaded applications, in which crWping is not a factor, a bearing outer ring may be retaived by a snap ring. The snap ring is installed in a groove ini the bearing housing, and bears against the face of the bearing outer ring. However, snap rings provide no circumferential retention for the bearing and, therefore, should not be used where the outer ring is rotating with respect to the radial load. For such applications, retention dips normally are employed. These clips are bolted to the bearing housing, and M't into slots that are machined into the face of the bearing outer ring. The dips thus can provide both axial and circumferential retention. However, bearing clips should not be used as the primary method for prevention of creeping. Proper selection of the interference fit will insuve that the boaring does not turn in the housing under load.
I
HOME
Spherical control bearings commonly ane retained in a -control rod by staking their outer rings. This method uses a circumfe-mntlal, V-shaped 5 coove machined into the bearing outar ring, as shown~ in the detail of Fig. 16.4. After the bewlng is pressed into the housing, a special tool is used to roll the portion of thes outer ring outboard of the groove over a chamber in the base of the holAsino. Thispoie positive axial retention of the bearing, and pemt the beoring to he removed from the asaumbly by te applying a heavy axial load. Because loading of_ bernsis primarily radial in direction, the staking procedure has been found to be satisfactory fbi most helicopter applications of sphc.ical control bearings.
16-4 ELECTRICAL FITTNGS GENERAL design of electrical systems is discussed in Chapter 7. Thin paragraph discusses the fittings used in these electrical systems. Eketaica fittings should be selected for reliability and eas of maintenauce. Therefore, seveal design pfincipae should be observed1. Mounting hardware should be connected per-mriPintly to the part bwaing munted. 2. Features should be provided 'to prevet incorrect assembly, 3. Right- and left-hand parts either should be identical or should be incapable of being interchanged, 4. Components should be protected against inCOrrect use of attachments. S. Universal mounting features shoiuld be incorporated where possibi!:. 6. The use of dissimilar metals in intimate contact MUbe avoided (see MIL-STD-889 and MIL-STD454. Requirement 16? for definitions of dissimilar metlsh). An interposing compatible material shall be 16-4.1
VThe 1¶
1
used if combinations must be axeombled.
7. The number of wire and cable junctions shouid be minimized, and only approved dcviccr &hal be used where junctions are required. S. The use of identical connectors in adjacent locations should be avoided, 9. Junctions should be accessible for inspection and maintenance. 10. Terminals and junctions should be spaced a sufficient distance apart to prevent arcing and detrimental current leakage between circuits. Because of the number and variety of switches, terWoonl blocks, connectors, terminals, and insulating materials available, manufacturers' data and the retwrened Military Specifications should be consulted for additional information not presented in this chapter. 16-16
Marking requirements of seleaWct lectical and electronic parts are defined in N4IL-STD-12S3. MILE-7000 contains the fameral requirements for sirborne electrical equipment. 16-4.2 CONNECTORS kND CABLE ADAPTERS Conniectors are used for Joining a cable to other cables or for joining zables to equipnmet in cases where frequent disconnect is required. The most cornonly used connectors have boen the circular and the rack..and-panel types. MIL-W:-506 does not state a eurmnfoaspcictcofoncobt roqueres that selectioin and use shall be in accordance with MIL-STD-1353. Although there are more than 30 Military Specifications dealing with connectors. the designer should reduce the quantity of connector varhiiions and limit the selection to those which lend themselves to common termination methods; i.e., common contacts, common bacd, hardware, and common assembly tools. This discusuioni is' imited to those connectors moat commonly used in htcicopiers. Table 1648 provides a ready reference to these typts. along with their general descriptions. Table 16-9 lists additional specifications to be used Whlcic Spa;' al trxjuzremenis exisi.(
Connectlor isa generic term used to dcnote both an electrical plug and a receptacle. A plug is a connector that normally is attached to a free-swinging electrical cable. A receptacle is a connector that normally i3 attached rigidly to, or is an integral part of, a
supporting surface. Each connect -r shal be selected to make tho "live" or "hot" side of the wrtnnoctor the socl'et member to minimize posible shorting when the junction is disconnected. The dead side of thec circuit i%the pin member. Thierefore, depending upon the individual cwircit aphig may mnta~~in ;he ins. or'sockets The mating receptacle, of course, will contain the opposite.
16-4.2.1 Connector Selectium A multitude of connector designs, with specific
capabilities, are available. Moreover, many of the features of these designs overlap. Therefore, the requiremen's pertinent to the titsk for which the connector isintended should be -ientified and then co.aipared with the features of those available. Information necessary for the selection of a,! proper connector includes the following: I. Spocific types of connect ors, if any, designatedi in the contract 2. Applicable enviionniental conditions 3. Maximum voltage and current for each circuit 4. Number of circuits to be accommodated plus spare contacts
9MW7120 TABLE 1"-.
MILITARY SPECIFICATIONS AND STANDAmiDS FOR CONNECTORS FOR AIRCRAFT
SPECIFICATIONS
I
STANDARDS
DESCRIPTION
SEE SPEC. SUPPS.
CONNECTORS, ELECTRICAL, TYPE.
YPE
MI-C505CIRCULAR MiL-C-5015"AN" MIL-C -26500
CIRCULAR
SEE SPEC. SUPPS.
CONNECTORS, GENERAL PURPOSE, ELECTRICAL, MINIATURE, CIRCULAR. ENVIRONMENT RESISTING ESTrABLISHED RELIABILflYY
MIL-C-38999
CIRCULAR
SEE SPEC. SUPPS.
CONNECTORS, ELECTRICAL CIRCULAR, MINIATURE, HIGH DENSITY, QUICK DISCONNECT, ENVIRONMENTAL 1 RESISTING REMOVABLE CRIMP TYPE COr4ACT RELIlABILITY
____________
MIL-C-83723
CiRCULAR
ASSURANCE P~dGRAM
________
SEE SPEC. SUPPS.
$PEC. SuP.
CONNECTORS, ELECTRICAL, CIRCULAR, ENVIRONMENT RESISTING, GENERAL SPECIFICATION FOR
MIL-C-Z4JW5
KcAUIArNU Pr~kEL Iot
MIL-C-26518
RACK AND PANEL
SEE SPEC. SUP.CONNECTORS, ELECTRICAL, IMINiATURE RACK ANO PANEL, ENVIRONMVNT RESISTING, 2009C AMBIENT TEMPERATURE
MIL-C-28748
RACK AND PANEL
SEE SPEC. SUPPS.
UUNEC~ttQIM,
CONNECTORS, ELECTRICAL, RECTANGULAR RACK AND PANEL SOLDEA TYPE AND CRIMP YYPE CONTACTS rICIMDAI
MII.-C-55544
I
FLAT FLEX CA'iLE
FEL SPEC. SUPPS.
I
I
PRINTED CIRCUIT
jI
__________
SEE SPEO. SUPPS. __________
5. T~ pc of attahment required 6. Wire ;ize, inatetial, construction, and other *
-characteristics
)7.
* 'J
\
crrcnicr,,ATAIrsA&
I ENVIROMMENt RESISTING,
CONNECTORS ELECTRiCAL, FOR USE WITH FLEXIBLE FLAT CONDUCTOR CABLI!, GENERAL SPECIFICATION FOR SEE MIL-0-55543
I MIL-C-55302
tLtU-LI KILRL,
RJ:CTANGULAR MINIATURE POLARIZED SHILL RACK AND PAEGENERAL 0PECIFISUP.CATION FOR
Type of couplifj required 8,Special mounting provis.oens req~fircd. The requirements of M1L-W-Stt) also should be wonidered. These cover safety-wiring of coupling
CONNECTOR, SOCKET, STRAIGHT THROUGH, FOR MULTILAYERED
PRINTED WIRING BOARDS
nuts, use of noninterchangeable connectors in ad*a cent locations, drainage provisions, insulation, adapters, aind scaling reqluirements. When coaditions permitz a choice, the crim-p style of attachment of wirc terminations is prcferred. The use of identical connectors in adjacent locations hall be avoided. Differtnce in size or insert arrangement is preforred. 16-17
706-;:-'FABLE 1649. OTHER MILITARY SPECIFRCATIONS AND STANDARDS FORl CONNECTrORS
SPE.$#IF ICAT IONS 4
'A'-
TYPE
MIL-C-10544
CIRCUAR
NL-C412520 22 9C MIL-C-22249 MiL-C-22599
CIRCULAR 1~ U A CIRCULAR CIRCULAR
MIL-C-24299 MIL-C-264821
CIRCULARI CIRCULARI
MIL-C-26599
CIRCULARI
MIL-C-27599 MIL-C-25169
PLUG (A) CIRCULAR
-
-CIRCUA
CIRCULAR
MIL-C-55181
CIRCULAR
MIL-C--55181
CIRCULAR(0MA;ANPNE
MIL-C-85114 MIL-C-815B2 MIL-C-81503
CIRCULAR CIRCULAR CIRCULAR
()RC
N AE
/
MIL-C-21617
RACK AND _______IPANEL
')F TCBEmnmr VflL W
Whiav identical connectors are used in adjacent locations. wires and cahles shall be so routcd and supported that improper counnections canrot be made. Adjacent connectors using the sanie insert arrangenmeni shall be selected to take advantKge of alternate insert po~itions or alternate shell keyi% pn3itions. If tiis requirarnent cannot be met, color coded sleseves having the identification o!the associated receptacles sllbe attached to the wires or cables nzar the plugs. The recptacles shall be color coded by a colored band on o- mounting structurc. 16.4.2.2 (liedir CoamIor Circular connectors are the 'most popalaa st: used for general aircraft wiring. This type of connh tor is shown in Fig. 16-5(A), and Table 1649 contains a listing of these connectors and their descriptions. .
16-18
RECEPTACLE
1-S
Flgure 16-5. Comome
IW
I tillt
lr11
tIW V.,vgISrI
Types of Ceustectois
Standard MS connectors arc available with from I to more than 100 c, ntacts. and in 15 different inscrt diameters ranging from 0.250 to 2.550 in. Connector si4e are based on the diameter of the receptacle shell, stated in sixteenths of at, inch. Standard contact sizes available range from 0) through 22. Either solder- or crimp-style terminations may be obtained; however. unless otherwise approved by the procurin&I activity, crimp-style terminations ehould be used. Standard pin arrangements for cylindriczi cornnecors conforming to MIL-C-5015 are centained in MS 336SU through NIS 3369(. Pin arrangements of other types of connectors can be found in pertinent MS standards.
)
16-4X.2.
Twuinadeal
Seal
Environment resistant cnetrhaigwe sealing grommets A be wsedwhnvrpsbe however, potting may be used when no connector having a scaling grommet is suitable for the application. Sealing materials should meet the applicable environmental requirements and sbou~d be salected from either MIL-S-8316 or MIL4-23586. 1641.L.22 Cable Adapter A cable clamp often is used at the end of the cir. cular connector to support the cable or wires and to prevent twistinez and pulling of the contact torminations. Avoiding this motion also helps to reduce the transmission of moisture along the wire. For applicable cable clamps and other ac&orroy hardware consalt the supplaement to thn,connector sped. ficailon. 1"112.23 c"eonctr C0*mPig Three basic styles of circular connector ccupinr ame used; threaded, bayonet, and push-pull. The bayonet and push-pull styles offer quick-diioonniect features. Other features, or a combination of fellturci. artz smonai'tmo used. For exAnIPIe, a pull ~to_ braaa)cinmyb noprtdwt h ai coulin cin a einop odwt tebs Pg. 16-42.3Rac an Panl Cmacersterminationsi. Rack and panel conkiector3 amc used to connect a cable toa fixed eceptadc, a cable to acable, or a moduet aTe acklat. comonconiguatin ~ rectangular or some variation of a rectangular shape (see Fig. 16-5(9) ). These connectors are aviitabl in many sizes and shaprs conforming to Military Sped-.
owl...
J
Refer to Tables 16-8 and 16.9 for listings of the -miitary types, and to the manufacturer' awj di*. tributers' catalogs for the cmmecial type.16.4T 16-4.2.A Flat Condeetor Cable Cemmetor Where flexible, (tat conductor cable i cued, can. nectors of the MIL-C-55544 type (Fig. 16-5(C)) should be considered. Thes# connectors ar: suitable for connecting flat cable to flat cable, to roundwlrs or to printed circuit boards. 164.2 5 P'lite WkIi 6 Boad CUU5scomI Priultod wiring boawV or printe circuit connectors (.Pig. 1&.5(D)) no~rmally arc not encountiired i.1 aireraft wir~ug. They an used fo- c:.nnecing priatod wiig -=azdo to conventionali wirikg, prinuid wiring \ boards to tchb ofhwer or jirinted wiring boards to a
r)
backplaw. Connectors in accordance with MUL530ampreforred frprinted wiri lato.
butg
16-43 TERMINALS Common typos of terminals currently in use indlude ~ug. eyele, and notched. Their Installation may be of the dip-solder, sotdeelem wrap, taper tab, taper socktt, or crimp style. Unieos othewise specified in the contractual documents or approveod by the procurio# activity. wire and cable tormitials &ouldbe of the crimp style. Crimp tormina1i, allow diu.ict contact of wirn ed turninal to be accomplse by deformstion. They can tio- intalled quickly with uniform and ui~able quality, even bý newly trained personel. The crimp can be accomplished wit hand took or with auto. matic equ-pment anid can be inpacted ceally. The crimp termirsas muy be used in hig-tmpwavure Wp plications whvm other types would ov unacceptable. li sireonnectHonwsevecric, m lseof n andm mediasisal-os Wsrn.Hwvr us fcipjit sams Impomsble. A,common practice is to porovidle extra wielntin whres satlssohta ere~~d1 e-0-etrnai' _ cutting otf the oid one and crimpirt on a new one. AMCP 706-125 provides a detailed die~wsin of the requirements and charatictiao of various types of MIL.STFD-195 defines tie mairklums of n fr elctirecMal03 term~bnal. rp co ppeti n S23 omnl Tj *,Frcpe i quiring conductor temp~eratures above 1051C. 17or suc aplctos uninsded tenninals (Typt I), conforming to MS 20659 aund to fthroquiruenats of MIL-T-7928, should be used. 17 ,r aluminumn cable, Table 16-10 containt additninal standards. MI L A S Whene wire ot cablu junctions require infrequent disconnection, or wherm it is necussary to join two or more wires or cables t* a comnsion point, teminal bowi4 shu be used. Terminal .boards shoild co~mfxm zo MS VnIZ an coer cofrmn to MS 18029 should be used with thm These boards lMaw mokldd barriers and molded-in studs. They may be purchasd in standard lengths and cut to the lengt (outabor of swds) required for each installtioum. Terminal board Iduttificatica should be in accordance with the para&Wah tidled "Junctions" in' MIL-W-5SM. lusme connecting the terminal studs should canform t,( NS 25126. 61
. 3
ltEA 106&
MULMMAY SPEWICAh1ONS AND STANDARDS FOR CUMMPJTIE TERMINALS
SPECIFICATIONS
TYPE
MIL-T-728I
MIL-T-7928
11
MIL-T-7099 MIL-T--21606
FERRULE
MIL-T-38732
SPLICE
STANDARDS
I
MS20659 h I0, 4 516 1 8 910 11, 12:1J, f4,'ANb 15
COPPER
1
MS17143, MS25036
COPPER
1
MS25435, 6,17,8 AND 9
ALUMINUM
1
MIMI121MS21980, M.S21 981 MS27429
SMELDED CABLE COPPER
164! ELCTICAL SAWITCIMS
0
146.1 GENERAL Military switches wwat be dWngoed to &btain mwdmtzmf~f Power capaciy withllL a limits sir and apace. Electrically, the moaut iriplortakit consjjaratlon Inv~ofre cotad toed radinj andl arrange amnt. Thesn factors iaLu gir~r 3ctAatz the ty" o& switch toOt tr"b In ci gl'w *ppfl"ain. Other cms&~ tching sequence inulderatlow nslded.detw sulrion vQWlsbt'au sl nt V~x fmftnoe, wvircnnmax asweds and any selot, Ictxarv that might be requidy. The koa-'aiynn rrqtumncos for each switch mint be analyvsd thorougbly. Stud. factors as v31t tody Ws cua'ct, sturg currunw, ftequency, contact best 6islpmlloit types of loads, and life requirewe tat eoiier&are The moat commonly. used switches are toggle, put.-button, slid*, @and rotary. Each one huo partimbe advantages which must be weighed for the *WgV tvae under comilderatica. MIL-STD-1132 mwtalMe sealcton van Installtion raquirmuents for
-4switches.
VWhere
wndonmight procr~o faaitch *=e serious wmwaqumaom, a switch guard Aid be insidled. Tins v*t$paW type of guard generally fall in.etok09 thse Axed or tbew hinged catgoyý F1SAh guardis are ehennel-ebaped, met almebers thgi require a finge to bv inserted law toe area
~W2
WIRE
CLASS
'r
between the cahannel-1 tag to acretuate thba swaitch. S c MS 24417 and MS 25221 for thins configuration. Hinged guards have a cover that swInsV down over the switch. This cover must be raised and rowaed out of the way bel'ore the switch can be actuated. Two types or hinged guards specfie in MIL-O-7703 art Type A switch guards, which h&'uc more than one maintained position, and Type R cwitdi guards, which are spring. loaded to thc clocer position. Military Standards dcocribing switch puard& inclu Ic MSZ5O'4.-25214. 25223, 25224, -25225, arid -25452. Table W6I11 lists snecificatijos and standards anplicable to the several typer of switches. 16-5.m. Tankl Switches Two types of toggle switch action are common: momentary and maintained. Those switches generally actuated by a toggle lever (ofte shaped in the form of a bat), and have a snap type of actioi. They usually have eitiier two or three positions: either side of the center and the center position. OtŽer arrangements are available as special configurations. The switching action may be one of many standard I.,"m, such as rocker-conact, positive-action toggle link, or others. Toggle switches are relatively low in cost, and arc used widely. PoshiattiS Switches 11.1 Pussh-button switches are dlusaified as momentaryaction, maintained-action, aid sequential-action
-
AMCP 706=
MILITARY SECIFICATIFONS AND STANDARDS FOR SWITCHES
TABLE 16.1. SPECIFICATION
TYPE
DESCRIPTION
MIL-S-3786
ROTARY
SEE SPECIFICATION SUPPLEMENTS
SWITCH ROTARY, GENERAL SPECIFICATION "LOW POWER AC DC FOR ELECTRONIC AND CbWMUNICATION EQUIPMENT
MIL-S-3350
TOGGLE
SEE SPEC. SUPPS.
SWITCH TOGGLE, GENERAL SPECIFICATION SEALED AND UNSEALED, AC AND DC
_
_
__
_
MIL-S-6743
PUSH BUTTON AND LIMIT
REPLACED BY MIL-S-8805
MIL-S-6744
PUSH BUTTON AND LIMIT
REPLACED BY MIL-S-8805
MIL-S-6745
TOGGLE ROTARYAN3212, AN3213,
"MIL-S-6746
*
IV.iEU,.UROTAR
._..
STANDARDS
**
__SINGLE,
MIL-S-6807
ROTARY
MS21994 AND 3 MS25002, MS90547
MIL-S-8805
SENSITIVE ANt) PUSH
SEE SPEC. SI PPS
MIL-S-8834
TOGGLE _
--IL-S-9419
I
_
TOGGLE
f
S~~MINt,!________R
MIL-S-55433
SWITCHES, ROTARY SHIELDED ELD I SHMI urii IfliHLlA r~,lTWO AND FOUR ENGINE SWITCH, ROTARY, SELECTOR POWER FOR USE INPOWER CIRCUITS SWITCHES AND SWITCH ASSEM. SENSITIVE AND PUSH (SNAP ACTION) GENERAL SPEC!F ICAl IOA SINGLE PHASE, AG, DC
MS21026 AND 7, MS24612, 13, A-14 14, 55 AND 6, MS25305, 7, AND 8, 10 AND 11
SWITCHES, TOGGLE, POSITIVE BREAK:• A ,4 CR ,F, '.T-,,••I~I~ ",, 4IL SPECIFICA rION FOR SEALED, AC, DC
MS28939, SEE PAR. 3 OF THE SPECIFICATION
SWITCH TOGGLE MOMENTARY, FOUR-POSITION bN, CENTER OFF ENVEk,O.'PE, FREE
_
_
_
_qlt
_
__••e\/•
RT E_ UR.N
_
DRY-REED
SEE SPEcC.JSU
I
IOPERATED,
SWTCHi CAPSULES, DRY-REED TYPE HERMETICALLY SEALED INGLASS MAY BE MAGNETICALLY FOR USE IN COMMUNICATION, ELECTRICAL, ELECTRONIC EQUIPtENT
l6.21
\.,
"",.
-.
types. A pushing motion in line with the buttoni normally is required; however, a modification of the man W-cintype requires a pull-to-operate Military ptsh-buttor. switch requirements are do. fined In MIL-S460. TM swtce descibe an available in five enclosure designs, four tinnpeatmr dwamaow"slc, two shock typos, and three vibration grades. The specification covers both sunsitive an pushtype switches. Sensitive switches are intended for a nonhand-operated mode. while manual (push) switches are intended for hand operaon. Th Supplement to MIL-S-885 contalis a list ef sped. ficastion sheets by switch title and a list V supersetled docuntents (MS. AN. and JAN standrds an specifications).
u
1".513 Reta" Switce Rotary switches are uted to control a numbee of circuits in consecustive steps. Thene awitches are available in single or multiple wafers. ahowing alarge variation in the configurations and making fth switches adaptable to many tasks. Detent: arn p-.oWid on. them switch= by th-. i=e of stee balls carried in a detent plate. The detents cams a requirement for a reatively h!&%torque to turn the knob, and the switching takes place with a snap a.-tion at maximum force and velocity. This action reduces the arcing time between contacts and promotes longer contact life. Most rotary selector switches are used in electronic ciruiltry wherm currents are low. The use of the sliding type of contacts on toese switches provides the necessar self-cleaning required for low contact resistance. Roiary- sw~tdma should be installed with the extreme countercockwise position as the OFF pouition. MIL-S-3786, MIL-S-6746, and MIL-G. 7090 dofine the requirment for rotary switches.
I"1"
PIPE AND TUBING FITTINGS
GENERAL Ai byuke stalled In accordance with MIL-H-SW4 unless otherwise directed by the procuring activity. Hydraatc "emt~lrquirm f~tU are discussed in Chapter 9 mmd the requirements for fuel and lubrication &yaomns ame described in Chapter 3.This paragraph discusses the fittings used in the installation of hydraulic systems and in the intaltio nof fuel and lubrication systems. Since each joint represents a possible leakage and trouble spot, the installation of pil4ng sytemis must be accomplished carefuly. The number of connectiona used should be reduced to a mininum in order to assure maximum safety and system ef flciecny with minimum installation and maintenance MOat. 164.2 TYPES OF FfI'FIGS Many types of pipe and tube fittings are available. For murcraft systems, those generally used are tapered pipe thread, straight thread using agasket seal, flared tutie, fi-lues t-W-W, thi-wail1 trftM, cui:.ct discoanec= and ponvuanent. Solder-type tube fittings conforming to MIL-F.400 ame inactive for new design, except for oxygen systems and engine primer lines. 164a2. Tapered Pipe Tresad Taperee pipe thm~ad fittings provide reasonably teakproof connections, but should be used only for permanent attachments or closures. IFig. 16" illustrates this type of fitting and the related specifications bud standard3. High-quality workmenship and machining to dlose toleances will hnurm a good seal with tWi type of fitting. Howevcrz Kcannot be used for directional adjustmems because it must be tightened properly to prevent leaks. and thert is a
TýPERWFITTING
EFIMNS. MI;BOSS.
PdIL-P-7 105
L
ms3p7
AND106
Figure 1646. Tapered Pipe Thead Flttlags 16-22
K
(
:
.
da& ofdamagingtheoomponent housnlgifthefltting is overtightened. This type of fitting is inactive for now aircraft design, except for oxybei systems, and should be avoided where psdble 164.= Stralsht Thud Fttlp Straight thread type are preferred for installation of fittings into the boass of a component, as lllutrated in Fig. 16-7. Standard internal thread dimansions for bosm are shown in MS 33649. The tubing end fittings should conform to MS 33514 and MS 33515 (flarlem tube stye); or MS 33656 and MS 33657 (flared tube style); or MS '4385 and MS 24386 (fisted tube precision style). Istallation instructions are givm in MS 33566 and AND 10064. Straight thread fitting instalations arm usable for nominal operating pressures up to and including 3000 pid but are inactive for hydraulic systems in new aircraft design. For Zhe universal elbow fitting installa^mrefer to AND 10000. MS 287• or MS 29512 gaskets should be used f-r producing the eal btwemn straight thread fittings and omponents. As the premure increases, the gsikt.,ges the oeaing eeac. This style of fitting rquires less torque to insure a goad mal than is recesary for pipe threads, and it may be disasembied and ressembhld without do. terkoation of the joint quality. Choice of the proper thread style (MS 33566 and AND IWO6) allows the fitful to be positioned without Impairing the scaling propertles. Installations of this type of fitting should be made "carfullyin order to minimize service problems. Ex-
*
"I MIL--772 MUS5 MS24305
16.2.3 FlnMd Tdo Fttdn Flared tube fittings are of the expanding type. The end of the tube is flared to an angle of 37 dog in tie mating slev. The slev and flared tube end then am clamped between the fitting and the out to form the fluid seal. Fittinp of this type are shown in F!. 16-8 and should conform to MIL.F-5509. Tho styles of fittings available are listed In AND 10019. The external male fitting end sheaild ciform to the dimenions listed In MS 3365, or MS 33657 for flared tube vonnections. or to MS 7A385 a MS 243•6 for the precision type of flared tube connections. For 3000-psi systems, components smller than 0-5-in. iu.w, maid u.of c. ..... Aluminum alloy fittingP should be used with caution in installations where topeated dissmembly and toasembly could damage the thrmead ftft end. The tubing end should be flared in accordance with MS 33584, except for aium!n.um &by tubing of 03CS in. or les (outside diameter); this, instead, should be double flared in aocordance with MS 33583. Them fittings are intended for ow in aircraft hydraulic and pneumatic systems. This type of fltting is
-FITTING
FL~eDUSE MS33057M,3646
MS24er I
trane caution should be teed when a gaket k instAll•d on the fitting. The gasket should rt be pushed over the threads or other am that might nick the gasket surface. A plktic or imetl thtadprotector or "thimblW" should be usad during Istallation of the gasket. Specifications for Installstion, gasket selction and lubrication, fitting torque, and positioning should be in accordance with MS 33566 and AND 10064.
DIMENISIONSI
FLARED 7U1E PRECISION
UNIVERS0A0
N
,4••
,FIg
r 16-7.
BOSS
SPACING AND1007 4
G__KETS1
Strallgh Thread Flultp 16-23
nu:w r.....
(.s
,
......p
is ma•. The fRnhe flae AW be inspwmed for pok marks, splits, Inadequate squareness or conwitricity, and other defects. AND 100M tefies avceptable value. for nut torque for swomNy of th joint.
inadive for hyI olk.; =:,,,ran ,r•fw airm-ft desg. Ifaflaridtwut oou mdionU uWdeimpropely. thMjoiut iuAý, ý,k. Mlvwwo, thu, pressur is not likely to cawm the tubes blow out of the joint. For ingallaa• n infonnation conowuin flared tubc fitfinags oft AND 10064.
7h1 tubing layout and routing shh•ud be designed inacoranewith MIL-115440 or MiL-P-SS It. The tubing should be fonmad so that no ta is placed uWon the fitting whce ft connetion is tightened. Suficient mruliht length should be allowe betwee the ead of a bend and the flarod portion of the tube for inatalltion of tho nut and aleeve. Tube cutoff must be sqmv, and ll burfs rwust bt removed before dw nut ad slew am pAed on the tube and the flre
164.24 FRnd. Tub. Fistap Flare4a tube fittinq art of the comp heion type, as shown in Fig. 16-9. Thv tube don not require a flinS operation. As the nut is tightened, the sleeve is compressed by the fitting body and deflects into the outer surface of the tubing wall to create an interfence se• betwen sleeve and fittin; and sleee and tube.
,_
'-•
FTrINS ,IL--,-•9 !FITTINGS•
•,
FLARED TUXE
S•
M=•
!
i
T M,•S4
SNGL
ITTN
MILA
;
PIECIS'ON
Mi24486_
THRAD
M L--S-774
Figure 1648. TING ENDS MS33514S3
2LJ
ITUBE
F~ared Tube Fittings TBN
FITTING
FLAELESS
MIL-F-18280
SLSEE TUBETUBE
INSTALLATON MS335F6A
Figar 16-24
MI
16-8. Farem Tte Fittanp
Fittings of this type are intended for wme in tewonautical fluid systems in accordances with MIL-H544 and NIL.-P-SS 13. They shoud conform to NIL.F.18230 and the instalation instruetloin of MS 33566 should be follow-d. The externO nmats fittig sod for flardeim tubing shotild conform to tde dlmsendouas of MS 33514 or MS 3351IS. For 3000-psi systems, tubing smaller than 0.5-in. used on the pressure or actuating parts at th ofOW. lumnum r seel circit beawmihoul may be used for return lines. Tubes measuring more than 0.5-in. may be of steel or aluminum alloy. Whomer repeatediassebl mW damag the threaded fitting end caution must be used if the finting Ws of aluminum. The tubing layout and installation should be do signed so that no bending or springing of the tube is required for aligunmet with the fitting. For flardele tube flitings. a tube of less-ductile tlfiawal tuing sm asbW~wgtb matmal --suha hlh-et~qt, tin-wll ubig mayen userd.s itnear m&a Whenst lruted s Routmay ans thwedo ben imrpel of the fitting when peuar eIsappied to the line. This ~ ~ ~ A fluid. The tubeshoul beecusquus &audWmnal and external bumn should be removed. Tha smt and sleev must be slipped on ths tube and the tWbn the. hel aanst the shoulder of the fleting while the Mut amd sleev are being dgighteed. TU a should be torquod ;n accordum. with the insetrimoasf MS 33565.
Fed hose coupling mequirwmaets ona give inNIL-H-
Cumintews Installations involving aimp volumes of flid suc as pneumatic systems or poe phant fa sytes should use rclati-ely WagM thin-wr.U tubing; standard tubes and tube fittings could result in unacceptably
fkrxeless fittings and can result in substantial weight
-
-
164..
TIwo was TWOe
high weights and pressure looms. Fuel systems shail be designed to the requirements of MIL-F-38363. For fuel systams where operating pressures to 125
psi are present, couplings conforming to MIl"-
(
7061 and NIL-fl-MMG.
63 o e-lnd V11011 ' Ql iesstC w Fofdadoilneat tcshofqck disconnect couplings should he used and should meet the requiremeints of MIL-C-7413 for Type I fael line couplings or Type 11 oil line couplingsi. Relleraw i shu emd oteseiiainfrdss Frtehdalcsl-eln.qikdsonc th yrulc ef-lnuc-isonc Forsebl couplings used in Type I and Type 11 hydraulic are covered by NIL-C. systems, detal requireaents 247Todse fculnsaedsgae:Ca
600 with arated pressure of 600 psi, furnished only in ae 12-n ue ieadCas300wt 30Owharte anCl tbsh .5i praiure of 30D0 psi, furnished in 0.25- through 1.0in. sizes. For protection against poet-crash firm, all flammable fluid systems AaM include &utoinatic~eutoff. breakaway fittings. These fittings zAe be designed to break away at the specified crash load factor. and to shut off and thus Prevent spillag of flammable Capt3 fliwds. Such culuputlems Oci dic ird in AMCP 706-201, and in Ref. 12, and in W14A1rD
J 164.2. Paesimnme Fktid MIL-H-544 requires the "ae of permanently Joined tube fittings except for production breakts and component removal. Permanent connections pro-.vide a higher degree of reliability than either fare or savings for the system. The four methoods in current t11' are swaging, braxins, welding, and cryogeic However, standards do not exist for these cam-. porints. Urage, installation, procedures, and test qienet amepmpriea.
16-7
CONTROL PULLEYS
2263 should be considered. Type I couplings connect two tubes with MIS 33660 Type A rolled bead ends, Type 11 cou~plings connect two tubes with MS 33658 machinixd fitting ends, two MS 33660 Type A rolled bead ends, or one MIS 33658 machined and one MS 33660 Type A rolled bead; and Type III couplings connect two straight end tubes. Standard Jcoupling sizes are from 2.00 to 8.00 in. This type of allows expansion and contraction, small
16-7.1 GENERAL Selection of the proper pulley for a cable system is important in achieving long cable life. Some of the earliest control systems for aircraft wer of the cablec type, using pulleys when a directional change was required. These systems gave lon&, trouble-hae service when properly installed. MIL-F-9490 requires that approved Military
angular and radial misalignmenwu. and smooth flow.
Standard pulleys in accordance with MIL-P-7034 be
-'coupling
[
used in 11ight control systems. The latter spea& usation cover the requinimenta for single-groove.
the bond between the pulley and its besihag. MIL-P-
saftifion-beauing pulleys of two typms and thre clasex T'ype I-Nonmetallt and Type lI-Mea&Uic;
miation on the strengths of standard pulleys.
and Clan I1Secondary control, Clan 2-Flight control, and Class 3-Heavy-duty control. For standard cuefiural ansMS 2D219, MNS 20220, MS 20221, and MS 24566. Performance and strength require-. UIUis and other data cr. given by MIL-P-7034. W6732 PULLKY SELKCFON The seletion or design of the proper pule fo aehieving optimum performance should be based upondevralguldlina. puley f te lages f~ ble diameter should be used, and the groove radius and the pulley strength "klbe appropriate for the diu of cable being used. 16-7.2.? Masy Damester The diameter of the pulley has a primary influeceio upon cable life. An increase in cable life of 1015 them amn be achieved by doubling th~e sine of a pulley from an initial pulley-to-cable-disotr ratio of less than eisht. Smaller. but still significant. ianUIS..I Provoneii aft UVeai " u10 the diainctOr ratio Incromse. However, increasing the pulley diantte resuits in Incaeased weight and space requirements, so the improved cable life must be evaluated against thene factors. Neveirtheless pulleys of les than 30 times the cable diameter should not he used under normal circumsftances; thoue with a larger diamete ratlo should be used wlker possible. 16-7.L23 PWIhy Gres" The radius of the pulls groove also is important, Uisa c~ahl in a nufllv where the Proove radius is too small causes a wedging actloit, possibty resulting in distortion of the cable. When the groove radius is too largek insufficlient contact exists, resulting in deformation of the groove tread and distortion of the cable, and. hience. the possibility of premature failure, Tb. radius of the pulley should equal onme-hlf the cable diameter plus approximately 0.015 to 0.030 in. for cable diameters to 0.383 in.. and 0.0M in. for cable diameter to 0.5 in. The contact between the cable and the pulley groove should be equal to approimately onethird the cable circumference. 16.7.33 PA" Sresipquired The rtength of a pulley may be limited in several ways dqepsodlng upon the material from wbich it ho maide. Coseideration must be given to 'be buckling or sltftin strength of the sheave, the checking and shearing strength of the flange, and the strength of 16-26
7034 and the MS standards provide additional infor16-7.241 Palsy Perflerssoaac Sufficent wrap angle of the cable on ft sheave shouald be provided to overcome the static friction torque of the bearing. Very =mail wrap angles should be avoided in order to prevent sliding between the cable and the pulley instead of pulley rotation. Standard pulleys use bell bearings that are groe-. lubricated and seale. The seal arn capable of witht~lfPSmtun* of -55? to 1216C. The sta~nd1 bearings are Installed In the pulley, and the assembly then is checked for wobble and wocentrilcty withln the liit $ ledWI MIL-P-7034. The dusog of brackets and the placement of guards should allow clearance for these wobble variations to avoid rubbing or other interference with the pulley mounting bakt 16-7.2.S Neameal~c Pleys Nonmetallic pulleys (Type I per b IL-l'-7034) should be fabricated of a material that meets the non-
tmawf~a~n-. In addi.... ,.., tion thsmtra-hudmoterqieet o
ftion tils materia should mette nocrroiequo irmn- or n5O5lestiSshudenocrav otn-r zin-coated carbon steel cobles, and should moest the other qualification tests of the specification. 16-7-3 PULLEY INSTALLATION Pulle installations Awel he dosigned! to insure that the cable alignment (the angks of direction of chang with the plane of the pulley) does niot eusced 2 do&. as shown in Fig. 16-110(A)ý The effect of cable sagtiog, wherm long runs ar* used, should be concidered when detemining the cable position. If necessary. fairlead and rubstrips should be used near the pulley to limit this mnisalignmenst. The desig of the pulley mountlnS ýzackst should be soch that the deflection of the pulley under load dam not result in a miaalignment angle In aexces of the 2-deg maximum. The slack-cable side of a system under load shul be considered to Insure that sagging, binding, an excmmave misalignment do not occur. The design of the pulley brackets, shol be such that there as room to thread the ca"l sds without the remo~val of the pulley. Spacers dha may be rw be'tevelo the pully, and the mouaitln bon should be made integral with the bracket wherever possible. To the maximum possible entent, thes use OfN thin shims or washers between the pulley and the sades of the brackes shoul be avoWde. Thus itmsin are difficult to in in and are lost or omitte usoft
AMCP puley (Fig. 16-10(9)). For very small wrap anglse, where the use of two guard pins is Imnprntiak, one pin may be used (Fi. -16-10(C) ). Whe two or more pins ar used, they r ity be offse from the point of tangency in order to obtain the space necuury for
CABPL 2d"
"MAND PLANE OF PULLCY
UW• CA (A) MAJJSM MISMLIGNMENT
GUARD PEIAT CAILE TAWGWiCYPO*IN I
2 GMe
SINGL•WARD PIN/
Sflactions, 7
1U
GUA• PINS , AT CAB.L TAN,,CY POINTS
1C)
GUA PINS ICM &ML CA.Es WW ANWAES
their installation (Fig. 16-10(D)). Where the wrap angle is more than 90 dog, intermediate guard pins should be installed (Fig. 16-10(E)).
To avoid possible bindnig due to relative do-
the support for guard pin should come from the same bracket that supports the pulley. The gap botween the guard and tlh pulley shuld be as small as possible, yet sufficient to allow for the tolerance variations, Including wobble and acoentficities. The gap should not be so large to pemit the cable to become wedged between the pulley flang
and the guard. The recommended maximum gap is one-half the cable diameter.
MAINTAJIN CAMILE AND PU.LE
CLEAIWMCE
LS...Ciuta
Spring guards should not be used in primary flight control systems. If used in secondary control systems, some method of retention in addition to their own spring effect should be considered. If used, spun
ONE *ImaIW ATE GUAM PIN
:
pins should be installed in accordance with Mb 33547.
OFF:S91 FROM*, TANGENCYPOINT
"16- PUSH-PULL CONTROLS AND (Di
CABLE GQJ o•S OcFElVM Poi ,NSCOW POIEN
)•
Fllgre 16-10. CAWle Allgammt and Pu~y Guard Locamm durinrl routine msintenont e thrir omicion could ca--e the pull"y to be misalignedl or induce undirable strese on the brackrt which would result in premature failure. fle anmdth c loading of flatmoentin tension in wesld, flancki should be of mountina avd the bend.no avoided. Moreovre, adequate back-up structure shoudd be used to provide a rigid support of the bracket bae For long cable life, pulleys in the same cable run should not be installed ctoser together than the maximum cable travel. In this way, no portion of the cable will pan over more than one pulley. In any cae, the pulley should be arranpd and located so that no pertion of the cable is made to reveau direction in beadIng when passing ovea two or more pulleys. 16-7A PULLEY GUAROS '-"'
FLEXIBLE SHAFTS
,,
Guards or guard pina should be installed at the approximate point of tanency of the cable to the
116-.1
GENERAL
Although push-pull control and xlble shdaft are similar in appearance. they differ considerably in construction and applcatl•on. basically. both devices consist of a flexible are that operate inside a casin. The cr.inh supposh and acts r a beai msedans for the core. The pbyt uil control is usW to ofansmit rote. linear motion by tenion or omprsitnw ofe or Tng ibe shaft is usd to tacum it powee o rotay motion, esually hoe curved psth, between two components. Theme two dievicdl ate desind specifically to perform their InirA dWW functions, and they should not be intemhanged. 164L2 PUSH-PULL CONTROLS The baiic components of the push-il control an the inner core (usually flexible), the outer tubular casing or conduit (rigid or flexible), and the end fitting, as required. The requirements for this type of control wre defined in MIL.C-7938. Pwh-pull controb am designated by grades. Grade A controb are made of specially selected components that are Individually tested in order to insure that backlash and operating forces are reduced to the minimum value. Orade 9 16-27
controbs are mads to hilg~uallty commercial standafds for applications not reqturiel' Grade A controls. Many types of 0rad 8 controils ano available with a variety of characteristics and style of end fittings to meet spewa requIrmnents. Applicable date are pro. vkide in manufacturers' catalogs, 7Ue advantags of the push-pull type of controls Include: I. Ease of design and installation because they are readily routed around obstruction 2. Lower weight and space requiremenits than pulleys, bell cranksi, and brackets 3. Corromion-resistace and permanenitly lubr5cated construction, resulting in ease of maintenance 4. A variety of end fittingsthat adapt readily to desire control configurations S. Ready accommodation of motion between the Input and output anchor points 6. Low cost. However, for a satisfactory desgnr using push-pull cONtrols, the limitations Of thi type Of syvstemn must be evaluated and diutenunindt to be acceptable. Such consideration should include: 1. Lost motion 2.
my-
f6w,
3. D* assembly'requiredtolinspect the siding core 4. Temperature llutltativas, particularly for asinnblles using nonmetallic seals, S. TINS 110008111 for the Outer housing to be anhoe rigidly act only at the end fittings but at regnia Wintuvals sloes its length (to control friction dad lost Motion) C Pow" weighst disadantage for long control ~ ampsabe eficiecy.The ~ (~d ~control 1~a emien" avraw"ba (strokC) from a pardrw lo@014 mPeol-dped upon its design anti does#al g mid. as1f so sid fittings arm attached oo domw em, the Woke is dependent simply Wepedo @muma ofam that extends past the fixed *mW 1111N Manly StYles Of end fittings amc avin somssescI configurations that wili acemmdm ccntrol strokes of frow I to 5 in., and mny be easto. tailored, 1442. CO&We leA" The allowable push load that may be imposed upon a push-pull control is dependent upcun the maximum stroke of the control. As the stroke incensue the eclumn length of the unsupported end also increases. and acorresponding reduction in compression loading is required. The allowable pull load usually is equal to the maximum ruWe capacity of the 16-28
control. and is independsnt of the stroke. As the number and sharpness of the bends and the length of the contro increase, so will the Internal triotdon. To awhiem the desired output loed, the input load must Increase accordingly to compensate for these friction .osese. This factor musm be considered when selecting a type or control. The possible increase in load. due to an increase in friction, that may occur during the life of the equipment also should be considered to Insur continued satisfactory operation. 1648.23 Core Ceulguadem The inner core sliding member may have a number of configurations. The simplest form is a single member or high -tensile-strength spring wire. This type of control generally is used for light loads, and in installations having a small number of bends of gecrous radii (Fig. 16-11 (A)). Cores of more sophisticated construction should be used for controls requiring higher load capab~litiua, more flexibility, and minimum backlash. Thse may be single or multiple strands of wire wrapped with outer armor to provide stronig but L~ta nmko, n
WI.P iC
fUi~fhdw. ,OV
Mnai%
of copnents that'include an action element, such as a thin, flexible member supported by balls that roll or slide iii raceways inside the conduit (Fig. 1611(C)). 'Thsen units may be custom-assembled for minimum friction losse and backlash, and high loadcarrying capabilities. 1642.4 Coesdmlt outer tubular casing or conduit of a push-pull may be either flexible or rigid and generally is clamped to the basic structure at frequent intervals. It must bt anchored firmly at the ends in order to achieve the desired control motion. Complete assemblies may consist of sections of both rigid and flexible conduit, The flexible outer casing usually is a built-up member. Typical flexible construction contains an inner liner of hard steel or plastic to provide a good bearing surface for the core. Around this liner is & structure of outer windings to provide longitudinal compressive and tensilc streigth, and to maintain the itrired flexibility of the contral. This outer member is sealed in order to prevent moisture and other foreign matter from entering. This seal may be in the form of a plastic or rubber outer jacket, or of packing between the windings. Rigid conduit may be built up in a manner similar to the flexible conduit, exoept that it uses rigid-metal tubing as the structural member. This tubing is bent
Q
"BALL OUTEROuIER
1A)
ARIMR WRAPPED WIRE CORE CASING"
16)
SPRING WIRE CORE
SLIDING WIRE
IC)
ARMORED CORE
IGRDOV
(
/
ACTION 'clr
ELEMENT FLEXIBLE RIBBON
BALL RACE CONDuiT SPRING WIRE SLIDING MEMIER
SUPPORT
FLEXIBLE RIBBON. ACTION ELEMENT
TERMINAL)
SLIDING WIRE
CASING
SLIDINGSLIREING
WIRE MOUNT
SLIDING MEBBER
RIGID END
ItULKHEADL)l
-
-
f
STATIONARY MEMBER
POINT (RIGID END WITH SWIVELI
FLEXIBLE
KNOB (MANYMAN STYLESS
STATIONARY MEMBER (HEAD CONTROLC
OPTIONAL) (D)
Figure 16-11.
TYPICAL STANDARD TYPE END FITTINGS
Push-pull Cables and End Fittings
carefully to the desirod shape and fitted to the installation. 16-2.5 FAd FAttlmp End fittings for push-pull cont-ols are available in multitude of styles and sizes. Many configurations have a self-aligning capability, either by flexing the outer casing or by using a slider mounted in a swivel, A seal is incorporated into the end fitting to retain lubricants and to exclude foreign matter. Fig. 16SI1(D) illustrates some typical types of ends.
16-4.3
CASING
FLEXIBLE SHIAT'S
There are two basic types of flexible shafts. One is
used to transmit power, and the other is used for control of equipment. The construction of these two types is quite different, and they should not be interchanged. Power-drive shafts are coistructed to transmit the ma.imum feasible torque. They generally arm constucted to rotate only in one firection, and in small sizes that can be driven at continuous speeds of up to 20.000 rpm. The casing generally does not fit closely on the shaft, and therefore the unit can be diaassembled for lubrication and inspection. Remote-control shafts can be rotated in either direction. They are built to p~ovide ,maximum accuracy and gene.ally are operated at low speeds. 16-Z9
Since they normally do not require periodic lubrica-
rouive atmosphere for special material considers-
lion or inspection, the contruction does not allow ready disassembly. The main elements of P, flexible shaft are cuing, and fittings, and shaft end fittinp.
tions 7. Required sa'ety factor, possibility of shuck loading, and starting overloads. Some of the advantages of flexible power shafts over other types of torque-transmitting devices ineue I. Power to equipment can be transmitted at odd angles relative to misalignments the drive shaft.are accommodated 2. Installation
1".13.1
Terqee Capaty
The torque of a flexible shaft is reduced the minimum capacity bond radius of the shaft is reduced. The
flexible shaft undergoes the most severn type of flexu-
2 ntlainmslgmnsaeacmoae
pattwhileit ibentttyef- two whenhitun al arm phase reviersals occur during ch revolution. As the bend radius decreases, the cable strands also underSo mome rubbing against each other. The minimum bend radius for a given shaft is dependent upon shaft sie number of layers, types of matarial. desired life, and the required torque output.
eamigy, allowing more flexibility in the location of equipment. 3. Driven equipment need not remain in the same relative position to the driver during operation. 4. Torsional fluctuations can be absorbed. 5. Cost is relatively low. 6. Rotating elements arc enclosed, thereby eliminati.g a safety hazard.
1441.3.2 Flexible Power Shaft Flexible power shafts are designed to rotate in one direction only. The torque capacity of a power-drive core in the unwind direction may be only 50% of that in the windup direction. _ wi.b a-haft ¢oe.n of the same diamete may have saignifantly difterent charactertics. The core usually is made up of a single straight wire wrapped with additional Isarn of wire each of which Lý wound on the preceinj layer. Sucmssive layers alternate in
16-.33 Flexible Cox"o Shafts Flexible control shafts are designd• to minimize the overall deluction and thereby provide the required accuracy. TLey usually can be rotated in either direction at speeds of less than 100 rpm. Intermittent oeratio,, to 3000 rpm, with a low (lea than 5 min) duty cycle followed by adequate rest period&, may be acommodated. Thewe control amrimbles usually t-re designed with a casing that coscly fits the core. The end fittinp
pitch direction. Cores may vary ccording to the number of wires per layer, the number of layers of wire, the diameter of wires, thi. wire material, the spacing between the wires, sad the type of construction. These varablas in turn, affect such core characteriics as torque capacity, trayuverse stiffneo, minimum radius of curvature, efficiency, doWhen mi ming a power-drive shaft, tle following
may be attached permanently. They usually do not require lubrication. If required, cores can be designed so that there is nearly equal deflection in either direction of rotation. When selecting a flexible control, the following ftctors should be considered: I. Maximum torque to be transmitted 3. Shaft length, in.
factors must be cosdered: I. Maximum torque that the shaft must transmit 2. Operating speed ranaes: a. Normal, 1750 to 3600 rpm b, Special, to 10,000 rpm e. Special small size, to 20,000 rpm (Searing or other means should be used to increase the shaft speed and reduce the torque) 3. Maximum torque capacity of a possible shaft configuration and effect of operating radius 4. Direction of rotation, preferably such that the outer layer of wires tends to tighten 5. Normal da•lgn for 100 million cycles at rated speed and torque capacity 6. Standard rating conditions, and environmental conditions such as elevated temperature and cor-
4. Radius of smallest bend, in. 5. End connections for both shaft and casing 6. Size of core diameter. Requirements for remote-control flexible shafts with a steel core and casing arc presented in MIL-S3857. Units built to conform to this specification arm intended for use in either clockwise or counterclockwise directions. The load capacity and deflection characteristics differ, however, depending upon whether the operation is in the winding or unwinding direction for the outer lay"r ot'thc core. The requirements of this specification should be reviewed for applicibility to the reqairement at hand. For installations and configuratians requiring special characteristics, manufacturers' data should be consuited.
16-30
j
AMCP 7*-20 1-9 Ma'
CABLES AND W[RES ISTRUCTURAL)
16-9.3 7 YPES CF CABLE CONSTRUC'flON
Cable ismade by stranding many fine wires, which
16-9.1 GENEkAL asstruturl meber in Cabls bcuse Cabls beusemy myasstruturl meber in special applications. They are of light weig'at, and they may be made flexible for ease of stowing and handling. Their usm as tension members - for opemaing cortrols, slings. and hoists, and as part of machinery - has resulted in a number of types of construction. Data concerning the design and use of electrical cables and wires can be found in AMCP 706-125. 16-9.2 * *Th:s
..
'
PREFORMED WIRE SMlAN1) AND CABLE
When a wire is preformed, it is hoJ~caly formed into the shape that it will assutrne in the r~inihed cable. relieves the internal stmnaim of thr, were and increases the useful life of cakit thail i; ý,ubjezted to repeated bendirigs. Tho total svreas whit., in the sum of th.- internal itressts arid tht bendi4g stresse, it, Wrt~cW by tI~e amount oi the internal stresses. Otherr advantages af preformned wire cables ame: 1-1can be cut without m~ixing. 2. It is easier to handle, haw les tendtncy to loop or kink, and is more trataujlt. 3. !t can be u"e wishi swaged terminals. 4. It has little, or no tondency to rotate, and will run true over pulleys, helping to reduce wear of the pulleys, TABLE 16-12.
MILITARY SPLCIFICATIONS FORk CABLES
SPECIFICATION
TYPE
MIL-W-5693
TYPE 1I1X7
MIL-W-5693 MIL-W-5424
TYPE 11I X19--WITH WIRE CENTER.-----..--NONf-LEXIVI.E 307 - WiTHOUT A CORE------------ FLEXIBLE 707 - SIX OUTER STRANDS OF SEVEN WIRES AROUND A "ORE STRAND
IOF
SU
- -
-WITH
WIRE CENTER--------- NONFLEXIBLE
SEVEN WIRES ---------- FLEX!BLE
I MIL.-W-5424
)X9
can be of vitrious materials. Howowe. foir aitcraft applications, corrosion-resistant material, as tpezified in MIL-W-5424 or MIL-W.S693, should be. use. These specifitiations cover wire strand and cables oi maytpsasdcrbAiTbl162.C te mabny tpes, asLW-3A des rbdin Tbc y 16se2.Cae blspeMLW-3Aalomybue. Tems oua y.o icatcbei seven-sirand construction. Each ntrand consists ofha number of individual wi-a. It has been 6etermined emicay tisteofonrcinwhsi i hstm fcntutowt cprciyta vuter strands, supplies sufficient roundness to con. tact sheave g~rooves with enough outer surface to offord all of the bearing surface required, and provides eno~gh contact points to prreent abrasion from being concentrated on too few exposed suriaces. MJL-W-5424 -^eifiwes other properties of the diftercra typexs o! cable construction. " CALSE CnN 1-. AL EETO Whien aslocting the cable fox a specific application, all peft'nent factors must ht considiered. The requiremeikss musi be naiached urruiv &_ut~k tiw V4UriUW#r cabla ,,npertics. A typical comparisoi\ that mfight bt riade wojld balance the following: 1. Cable strength anid maximum load 2. Cable stretch &&-d alloviable deflection 3. Operatink c0arný&cistk~s and systan friction (if applicable) 4. Wire niateriul and environment 5. Cable constract ian life. and ab-m~ion.
-OF
7x19-SIX OUTER STRANDS OF 19 WIRES AROUND A CORE STRAND 19 WIRES A----------- FLEXIB LE f
MIL--544
WC SIX OUIER STRANDS OF
WRES ROUD A7x7INDEPENDENT WIRE ROPE CORE ---------- FLEXIBLEJ 16-31
#~
AVwires
16.4.1 cabe S&Iragth The breaking strenot of acable is determined by the nat ipatalic acrs-ectional area aipd its material pro~ws'Ie. The net metalhic cross-sectional area equals dhe Mmiof the cross-eectionsi area of all of the individual wires sand. therefore, varies, for a given nominal cable.
I. -detorrmined
~ Fother
The construction of the cable can take many form
164~.41 C"bl Defllectlea The total cable ddfiction may resul from either cosrctional stretch or elastic stretch. The maxinuni stretch allowad by MIL.W-5424 is2%wihen the cable is loaded to G~% of its breaking strmegth. Constructional stretch varies from cable to rable. and results from the small spaces prmaet between the and the stania, anJI between the strands an J the ccre aP=e fab-ication. Cables used -n cont-ol systems, and where large initial deflectiot. cannot be tolerated, are proof-Io2ded to remove this stretch. The caible- should Se loaded to a minimum of 60% of its breakiuF strength in accordance with MuL-C5668. This procedure also is useu for proof-testing cable assemnblies, Blatic stretch results froin the rlongnution of the inGiviGuxi wires as the iosa is aripiled. As the kaci is released, the wires reurn to their original length,
to obtain the reults desiasad, a decuauad in par. 16-8-3. The type of construction wead should be matched with the operating re~ulem tsufor the cable. A isystem might be designed using sesmatts of cablin with diffeirent construction. Aircraft cable measuring 7 by 19 (7 strands of 19 wires suxt) isproeferred for aircraft controls because of its high strength, good flexibility, and bendling fatigue meuistance - which allows it to be operated ovw rclatively small sheaves. The 7by7 (1 sirands of 7wires) aircraft cable isnot as flexible ssthe' by 19. Since each strandis made up of only seven wires, each wirt is larger in diameter. Therefore this type of construction has more ability to withstand abras.*on tnan does 7 by 'I.9 The I by 19 aircraft strand is considered non* flexible. It shouie be used only as a straight run section bemause its minimum stretch results in increased rigidity in contrql systems. It also isused for bracing. etc_, where its compact structure, hilth strength, and minimumý stretch all provide added benefits. Othein airciaft and commercial cable types ame ovailabie. Somne are aesigned for higp enerogy absorption - as in launching and arresting systems,
exceeded. The elastic stretch is determined from the. product of the loadi times the cable lm~gth divided by
coate rable, partic-ilarly desirsil~e for use as a hand. hold or guard rai!; and as comiplcx assemblies used
the product of the metallic CrOXAsCCtio~al area andl
for comniunicttions, where electrical conductors
the modulus of elasticity, 16-9.4.3oute
with insulatioc. are provided inside the loaci-carrying members.hrstehd
pmwided the elastic limit or t *ematerial h&3nsrot been
Sthe
164.4.5 CAibe C Mu oinle.
The operating characteristics of a cable are dep"IdnnImt upon the tyre of constriction and the knstellatlon. A firction-pmeventive compound nmay be applied during fabrication to reduce the friction of the cable when it is bent. The number 0t bends, the radius around which the cable i3 bent, and the pulley comifsiratfion in which the cable operates all affec-t open sting characteristics.
tow"Ir-,
-.
etc. Other configurations include plastic-
SAEYWR
AN
C0 a i EK riNS
?ins always should be made iafe. In main structural members, safety can be achievad by drilling a hole and using a cotter pin or safety wire. As an additional safety measure, bolts and pins should be inserted with their large end, or head, uppermost in 16-94,4 ~deMateialorder to reduce the possibility of their falling out should they not be properly safetied. The comptotition of wire used in the fabrication of In generail, safety wire should be used only where co,7rosion-maiiuant istocI wire rope isgiven in MIII-Wself-locking fasteners or cotter pins are- not adequate 5424 and MILW-W5693. The physical properties are to withstand the expected vibfation or stress. Safety by the manufacturer in o~der to mect the wire should beattached sothat icaberemoved i "reuireiaim~ts -f thc Military Specification. Many accordance with MS 20995. Inconel (uncoated) and types of materials are available, but their use Monel (uncoated) wire should be used fT'm all general shuuld be considered only with the approval of the lock-ivIring purposes. Copper wire that hrs bce!, Pnocurng activity. Oil-e. types of cable inclu~de galcadmium-plated and dyed yellow should be used for vanized carben steri (MIL-W-6940 and MIL-Wshear and seal wiring in ordcr to allow operation or 1511) and no! magiietic corrohion-rmsisting cables actuation of emergency devices. Aluminum alloy kMIL..C-t 8375).
16-32
K
ALCLAD 30S6, anedid and dyed blue in accordmmo with FP.STD-g, should be used ex.
dusively for nifst wiring of magnesium parts. Ali of
0-610MAt-TO
these matoriab can be Identified viasdly by teir
1U wire dsould measure 0.032 in. in diameter as a minimum for gewA purpoew The doubkb-twX method of w"ety-wirin dsould be conealaed a, standard. The singl-wim•e me0tod may be ued In a cloaey saced, dosed geometrical pattrn, on part in electrical qyumns, &nd in other aWlications that wake the single-wire method xmowe a:a.. Parta shel be safety. wired In such a mut" that the lock wire is put in tmnsiun wLen the part loomes. The lock wire always shouid be tvlatd w that the loop around the head stays down and does not coame up ove the bolt head, leaaing a slack loop. A pigtail of 02.5 to 0.5 in., or about 3 to t twistbahould be lei at tCe end of the wiring. The pirtall Oould be bst backward or under to prevent snagging. Fig. 16-12 U"lustrates various lock win applications. The two o•f safety wire and coter pins Al be In accordance with MS 33540.
v •'MI Id,
CLOW,c
, .'-•"
k OA
,
1>
.i-mw
"N ,
flg.
.M
RIg.bad Thred Tyir Appl1642. of Apoketw of Saf,' Wir
"REFEURNCES I. A FBMA Samdv*d, Section No. 9 - keAichod of Ewthoaing Load Ratins for Ball Bearlngr. Th. Anti-Friction BDuring Manufacturers Association, New York, NY. 2. AFEMA Stmarhs. Section No. II -- Method of Ewvmdnb Load Rdling. for kaller bearings,The Anti-Friction Bering Man-darturm Association, New York. NY. 3. T. A. Harrm, Rolling Hearing Aralysis, John Wiley and Sons, Inc., NY, 1966, pp. 221-241, 4. A. Palingren. Ball and Roller tearlng ring-,
i'"d
o
%P-.
,,, .
7. Aein GQu'sit Cwclag Wad E, reft MP i. Aetne Bali and Roller Bearing Company, Chicago, IL, 1960. 8. Totrington Bearixgs. Ctalog 567, The Torrington Company. Twrington, CT, 1967. 9. SKF Spherical Roller Bear ng, SKF Indawuurs, Inc., Philadelphia, PA, 1966. 10. P. B. Lindley, E&gneevft Pdulg with Nat".. Rubber. Tie Natural Rubber Producers' Remarch Atuociation, London. 1966. 11. 'Seals Reference Issue", Madabe Desrg. 43, WN 22.; qntembex 13. 1973.
Philadelphia, PA, 1959, pp. 115-134. 5. "Bearings Reference Issue", Machine Design. 42, No. 15, June 20, 1974. 6. Lubricativ, Guide, the Fpfnir Bearina Company, New Britain, CT, 1964.
12. USAAVLABS Technicst PReport 714, Quehi.worthy Fuel Sydtrm DeSxgn CriteriM and A•mWy ,u. March 197 1. 13. ASTM En~lneerin Desag Gddre. Life Adjutment Factorsfor DWI amd Roller BBewis.
13
:-
,.
\16-33
.....
K
-7
CH4APTER V7
PROCESSES 17-1 INTRODUCTION und in mie tM dcacap3 This ufacti"ars mad Jsembly phases of heliopter congeal. It W tho bc mtalwoakn and shuetprocaeres - camq, forging. extsj NNal f aWl. Machine shop Iectim and b d. 7U typsM of machine tooks also are mathods of jolaft -_ sch as wvldag, brazg, and oldrim - are dinutamd, as aft the processes of and cable , bonding swaginr Michani cafa spl•lS. Various typos of beat treatment are deM~bed, includin sutem relieving, conventional d suifece quenchiag and tempetin, aging. hardking. Work hardani techniques, such a shotL- addr.idng, with paminig and b the types of materials and Peus whe fatie VP_
U
t-i.
O procfis
Tft
tooaling rquliremet nosed to produce a #iven typV . or part ate re of bo Vendors shotld be used as addional iWfwation soureus bcacusN of the continuing development of t now ngerw and new poiess tb ht and performance relatiomhts between alteruative ro e tiufa~tui
17-2
METALWORKING
174,1 GENERAL This peragraph discusses metal-forming processes and their applications to tha wco ction of helicopum a wll as soei of the parameters governing the choice of ona pmo among many for the produetion of a particular part. More comprelhnive di suom,a well as dvJiled deip data, are found in other documents suc as AMCP 706-100, MILHDBK-5. -693, -694, -497, -6, and -723. Abe, Ref. I Is an Important sou for design data and maudlu*Wia details. Th, primary mete! frrication proceses are catig forging, eztnbh, and sbw-mnetul forming, The cbole of the appropriate procs depuda upon tlse s" and conapl y of the pan, the nature and manptds of the strium to which I will be subject, the material from which it will be made, and the relative coots of fabrication. In peneral, large, fairly coapmi perts - which require high rigidity, e not subject to exie:l stress or impact, and for
which somewhat rester weight can be toleratod can be produced more economically by casting. A part that will be subject to high arats and in which toughns also is required - and where high stregth, lighter weight, and better finish also are necossary - •icght be fotmn4 better by forging. Extruiion fonnirig may be preferable where highstren , doem4olerance parts lkaving a contiiuouis contour - e.g., rails, tubing. and beams - - are raiuired. Large-area, thin-wall, deei-draw configurations mote often will be formed from sheet metal. In most cans, some machining, joining, and Jmiahing will be involved. On'y a detailed analysis, with a careful evaluation of the several process and material co"ts, will provide a sound basi for selecting the memal-for.•ing process for a particular part. -
-n-,
andi*p4.n~~
parts will constitute the major portion of any helicopter. The parts produced by these processes range from the smallest and most precise instrumentation to the largest castings for gearbox housings. The materials employed in these processes include iron. stees, high-performarnce alloys, copper, magnesium, aluminum, and titanium.
1
CASTING
Metal castings are fomed by pouring molten metal into a prepared cavity and allowing it to solidify. The casting processes most commor.iy used in the manufacture of helicopter components are sand, in. vestment, permanent mold, and centrifugal. Casting is selected over alternative methods primarily on the basis of cost. The strength of castings generally is lower than that of wrought alloys. However, the structural properties of castings are the same in all directions, and, therefore, in case of symmetrical loadings a casting may be the most efficient design. Castings should not be mployed when the prev minant loadingp arm not steady, i.e., when tht loads either are alternating or involve impact, because catings do not have the toughness of wrought alloys. The possibility of the inclusion of sand or cther impurities, blow holes, or other invisible flaws results in the need for careful quality control of those castinp used in critical applications. MIL-C-6021 provides the standards for dassirication of castings on the basis of the hazard following their failure, and also 17-1
specifics applicabie quality control roqu.roemsn.
cs;lmitations onthe ihape. :lz and intricacy of
quirements are given in Chapter 9, AMCP 706-.103. 17-2.2.1 Samd Castdsp Satid casting consists o' forming a mold from san,I coicised around a suitable pattern of the part tw be made. removing the pattern, and pouring ~ molten metal to roplace tho pattern. The advantages &this piroc~a are: almost any metal may h~e used, thr-m i3 almost no limitation to the size or shape of the part, and the process is relatively low in cost. In adi4tion, extreme complexity is possible, tool costs sre low, and the process provides the most direct route from pattern to mold. There are also disadvantaps:tricate 1. Sand castings have rough surfaces, 2. Close tolerances are. quite difficult to achieve. 3. Some alloys develop defects. ~~~~4. Long, thin projections are n~ot practicable.
-,
if
~~~~S. Some machining usuAly is necessary. 10725331metal
Investment casting uses patterns of wax made in a split mold. These patterns can be combined into coinpV~x assemblies. The assembly then is "invested" by coaiting, first with a fine slurry of refractory powder andbinermaterial and then with progressively coarser layers of sand. When the resulting mold is baethe wax ismelted out (lost), leaving a smooth..
17,21A. Csalft~jigat Csutigp In sand and initestmcrit vustip.%, the mold is filled with tncizia simp1 by thc force tof rsvity. In centrigl~~dOftj~~'~~i ttua oc 07 imn by fpinnin tfh vi eMsuert Pth wonath t mci is bysint b'rs vAkteuoa w. s cagdt t h to~ee a erttdaa two axea. Ctll'rifu&-ý fhc - IP1 fitl theWA 1 pletely. Gases aamd impw' irle are ;oxatak0 near, teCne oain n s~ ccsaekp a minimum. This method is particularly *4&*rý&bk :to s-aaal, 'Incastinigs that otherwise watild be difficult io gate. A good surface finish is ebtah'te. However, tolgcoafrceriul atn 5 hhadte castings rmt. be symametricaal. Alloys of separable meusnynobe'stitdeel. p2 OGN Thc iorging ti *amcial pain ~invulit~ hswtidra- -a blank t~o a plastic state and hammering it into i-hape. In this process, groat ttrength it imported because of the bene~clal grain flow that takes place as a reult of the kneading action on the metal. Ther grain flow is changed to follow the contour of fthpart, rosuiting in a tough, fibrous structure. U;Maly, tht shaping takes place betweev closed dime that determine the contour of the forged part. ForginpS have
Thsprocess has the advantages of high dimninlaccuracy: excellent surface finish, and yuflhinifiin iafiuicacy. Moreover, any-s metal may be used. However, the size of the part that can be formoid is limited. the labor cost is high, and expensive patterns and molds are required.
~
,
~
17-2.2.3 Pernameut Mold Custlgs In the sand arnd investment casting prove.mes, the mold isdestroyed duving the removal of each casting. In permanent mold casting, the molds are madec of metal, usually cost iron, die steels, g.aphite, copper, or aluminum. The permanent mold is machined for dimensional tolerance and draft angles. Vent plugs are inserted into the cavity to allow gase to escape when the miolter. metal is poured. The process is readily automated, with mahy molds on a turntable using a fixod pouving, cooling, and ejection cycle. The ptocrs has the advantages Gf good surface finW- &kid grain struciure. high dimensional accura~y. rapid production- rate, low scrap rate, and low porusi~y. Dieidvatasges inulude high initial mold 11.2
sistance to fatigue, and are used for applicstivn.z isi shafts, axles, springs, ihairs, rotor hubs, rind similar moving parts. in the detiagn of forgings, it is AX sary to consider the lower level of strength normal to the grain flow. This cionsideration may nLfgvt thc orientation of the grain flow as well as the crci~aection of the part. The coist of foigings is subotantixly higher than that of caitinp.
1-.
XRSO
XRSO 1-. Extruding is a process in which a billet or slug of metal is pressed by a ram until the pressure iuiside the work piece reaches the flow state of the materisl. The material then i,! squeezed through a die thii cornkains an orifice &~the desired shape. Became uf the high m.duction ratio, the; ietal has excellent transvese fluw lines. This provides greater 5'trength in 'zbe longitudiiaal direction and lower strength in the trans,-qrse direction. The nonferrous alloys of aluminum, magmiesium, and copper are used most oomnmoni~y for oxtrusions. bu some steel alloyn are extrudable. Very complex shapes are possible at a cost much
LI
t*.9
: ~* ,
.
3m. then tha of uwcining. Extruded shape often can replawe %f-lt and puts previotsly madJm1d from bar sock. Te cost of extrusion die is rdadvely low, making &Aort runs practicable. ExtrUded trmis and Wkt baa, bas and tubes, and stringr find rady c•p* atlon in helicopter structures, 17-,.
SHEET-M"ETAL FORMING
In addition to lr. sarfces -real. such a fuselage sbells and s"us, a mutitude of smaller parts are made by set,,mesal forwing.These include stralps, brackets, clamps. cups, presmare housings, headers, Insrs gromnmets strinrs, and conduits. Sheet thicknms normally is 0.013 to 0.141 in., but thicker blanks may be used for vanous draw-forming oPertbies She matbri••us fror which metal p-rts for h i are tormed include alumlnum, titanium, helicop stersare fo dnusually
17-241 Ma- m. Fenrlng Many sheet-mn parts are made by machine#-- n ao , ftian6 at sart with sheet-metal blanks. The operations take place within dies, which are moowfod on various machine designed to supply ,_•Cti:& forces io shape thi materials in the dks. There are dies for cutting. bending, squeezing, and drawing. The material may be subjected to a combination of sw al, or all. of these operatons in s single die or in Ssuccssion of dies. Certain operations are named for tO manner in which the force is applied, By far the largest amount of cutting, bending, squeming. and shallow drSwing is done with the drop h..imater. In this machine, a failing or powered weight .. oucc = ., .. l_to nerfnnn thefrneinsary function. "In hydroforming. hydraulic pressure is furnished by a fluid to one side of a rubber diaphragm. This "diaphragmpresses against the metal blank and forms it around a punch that moves up into the forming mold cavity. The pressure causms the metal to flow evenly around the punch with little thinning or s;ipping. Hydroforming usually can adhive in one or two operations the same results that would require four or fiv, draws in a normaal press method. It usually is employed with higher strength alloys, In stretch-forming, the sheet metal is pufled ovti a form block. The material is strethd beyond its cln*tic limit, causing it to take a permanent set in the desired contour. There is no spring-bo..k, but allows= must be made for dimensional cha 1gas in the meta. Large parts with compound curvatures such s the external surfaces of fuselages, cowliugs, and fairings - may be made in this manner.
...
.
.. ..
Spinning is uas. when cold forming circular, symmetrical shedt metal paWi such as pans. covers, shields, and bullet -ee shapes. In spinning, a flat circular blank is clamped on a die or chuck iii a lathe type machine. The blank is Kevolved or spun and the metal is formed over tin chuck, using hand held forming tools. Parts with return flanges ray . formed by using band held forming tools, or by using collapsible or take-apart spinning chucks. Parts "n be naimmnd to sia on the machine, using conventional im• cut-off tools. Wooden chucks can be made economically for prototype and short run parts whereas aluminum and steel chucks are suitable for produion. In explosive, or high-energy, forming, shock waves ac generated oy explosives such as dynamite, exploding gSase )r an electrical discharge. The shock waves are transmitted through a liquid medium, water, to the work piece, fanning out in all directions and forcing the metal into a preformed die cavity. Materials of very high strength can be formed in this manner. The same materials formed bX other methods would have excessive spring-tbck, but, due
to the high energy rate and the uniformity of distributioa, very ittlc spint.-back c•cforming.
aftcr c.plosive
17-2.5.2 Shdp Fadcadem. Although the machine-forming procses previously discussed are used in the fab'ication of a large number of parts, the majority of the sheet-metal work involved in the manufacture of helicopters is performed in a sheet-metal shop. Here sheet metal is cut into various flatwork pattorns; punched, drilled, folded. seamed, crimped, beaded, grooved, turned, rolled, and burred; aid then joined by clinching, soldering, brazing, welding, rivetsa8, or adhesive bonding. Many design and fabrication techniques are available to add strength and rigidity to simple sheetmettl parts. For examplei strength can be incorporated into the structure by men of flanges, ribs, cor. rugations, beads, etc. These and similar metalworking operations can be performed either hot or cold. Hot-working may involve little or no strain hardening, whereas in cold-working, considerable dislocation and strain hardening can occur. Work hardening can produce beneficial effects, such as increases in tonsile and yield strength, but accompanyirk# dccreascs in ductlity and toughness also may result. ReSardlesw of whether the sheet metal ii, machineformed or worked in the shop wit'i press brakes or by hand, the metal will be subjected to bending in many 17-3
ways. The em with which a ntal cait be bent is de. pemdmt on ma'y factors. The Vame important ae the amn olf he metal itmelf, Its hardne, temper, and pWar working, and the orientation of the bend rota0"iv to the dic•tion in which On tbWhee was rolle. In Table 17-1 data are peme % to lilnetra the widc wrlety of characteristics avkiiam from working d for repreWar sheet natls Minimum red# ana estative metals In Fig. 17.1 omrnau fg operations, and their recomwmded bewd cur" iflustrated. radii 17-3 MACHINING 17-.1 GENERAL This paragraph diacume machining practiend
TABLE 17.1.
the relationship bewee the deigner and the machine shop in the contructimo of helicopter Pats. " a detal More omprebuisv dlmeen,m well design da, will be found in Chapter 1O, AMCP 706-100. and MIL-HDBK-5),-6934, 497, -64 , and -723. As distinguished from tbh forming opeatons diocused in the preceding paragraph, macining involve the removal of material from the work pice. Thus, proper deign, dimeoiinlng. and seqummncg of operations are important because an error, ome media is not my orected. Prope, design of the part is esmetial for a quality machining operation, and the design must be complete in every detail before any material enters the machine shop.
SEND CHARAC1ERISTICS OF SELECTED METALS
STEEL: TEMPER AND CONDITION t'.
-
t l.i-^U
..
BEND CHARACTERISTICS
!-I
.rn
%N, VU,
A@1•C oAnI
r - t PERPENDICULAR TQ ROLL
No. 2-1/2 HARD-RB 70 to 85 No. a-1/4 HARD-RB
(.
90 deg PARALLEL TO ROLL 180 de._PERPENDICULAR TO ROLL
No. 4No. 5-
180 deg FLAT ON ITSELF IN ANY DIRECTION
RB 65 Re 55
CARBON S 0.0i6 OR LESS CARBON %0.150 TO 0.25
180 dea FLA' ANY DIRECTION 180 deg f - t
THICKNESS t, in. _
_
MINMUM BEND RADIUS r FOR STEEL
_
_
MINIMUM 45,000
_OF
TO 1/16 1/16 toI/4 1/4 to 1/2 ALUMINUM ALLOY
,
YIELD STRENGTH psi
50,000
t/Z it 2t
It 2t 3t
MINIMUM BEND RADIUS Ih 1/A in. FOR THICKNESS tOir.. 0.025 0.032 0.040 0.050 0.063 0.090 0.125 0.250
TEMPER
0.016
0
0
0
0
0
0
0
ALCLAD 3004
H32
0
0
0
1
1
2
5254 AND 5254 AT TEMP.
H34 H36
1
1
1
2 2
2 3
3 4
H38
1
1 1
1 1 2
3
4
6
3004.5151
I
17-4
0 3
2 4
8 18
5
6
6 9
24 24
9
is
40
4
....
.
tZt
i'7~ -.. . muing chips to come off in sapuents. hi makes the
TUBING
`DRAN SHELLS
a
. .. .
a x 0J25 d
attainumt of a fnefnish diffcult.
Important fetors in the overall cmt of ma=hbinu
are dwwol m-aW, ___________d_
Mr2a0 finish tool temperature, and tool life. A major parameter is borsepower. Table 17-2 lisds the relative values of horsepower required for turning. "drilling, and milling repsentative metals. Horepower, surface finish. and too liUf may be used as a measure of machinability by compiling the rmults obtaine for auch with the results obtained u ider the 5 TIMES MINOU FOR TEN THE M METALALIUh R=3TO MINIMIUM BEND RADIUS
(B' 'LING 17-i.
Sam&"
Ma
ge smetry, and tle wat-
d.huabllsy of the metal being maa*iW. Methods of meauxrng the nia ebillity of a given matrial we bosed upon cutting ratio, shear anglei horsepour,
(A)FLANGES
IVnt
t
Ung iaW employed. A secondaty factoc it the war-
Pal
Ptdii
-
on" E.I
pm•,
s ihSA 11 sel a.d10 maId ot(Ia~ n machinabble are turntbree major m ,'b'n shopother operations operations arc and drilling. All ing, m!lling derived from thee three. In turning operations, the work plew is turne against a cutting tool. In mifling operations, the work piece remains at a fied height and is moved back and forth under an arbor or spindie hnidpin the mubibladiad cuttfin tnl. The tnal ri-
tatm on an axis paranhe to we plane of the work t7&J MAa NIWG OPRIATIONS The mc• ihre•'d to in this dtmiscuno ar cut-
which are dig tool ty•s Qew tyWp of aumai dpe%, include electrical discharM, msed to a liessr
table, and feod to the work piece is made by vertical adjustment of the tool. Drilling operations arm conducted by amas of a drill t-,Anlg turned into the work
pwec.
al, ukrumai•c, abrasive je ehtoemlcl•md c ad pma ar. All can be programmed from uamerll controi by tap ar by moputer so that only aominal supervision is required. Thia maults in
17.33 ELEMENTS OF MACHINING DESIGN The umoothnsse of the srface produced by machine-cutting or metalfinishing operations is expImed in RMS microinches. Measurement stand-
reduactoms for quatity and quality
ANSI by extvtine ards for them surfae textures re iven v l
Amowq tho operatdo performd with cutting tool t reUp1mnnuag planing. Aang, machiss an winig noutiand thread milling broAhn dilg sni ueWboq an die thradlig. A tool that cu with one point o"r odg Is rehnd to as a
ojertions are given in Table 17-3. The tolerance aid limits applied to-a dinunsion on the drawing of a pat will determine the kind of machining proess that should be used. Thus, for a dimemion requiring a tolerance of 0.002 in., it is ap-
and millin cuttes
prpriate to use a grinder rather than a lathe or
Was. hayv mo thUm oas cut During machiing opertloai, meta removal taIes plea tkmhrog threm distnc types of cutting action, speedbl upon the type of material being cut. MMim•ard materiahl with a low codflciet of ktion t oww a coatitios chip that ten& to foul whle otW ad ductaie mar& with a do UWek, high o0feledat of ffiction give a continuous chip
milling mcbiuine sce the time spent trying to hold this tolerance on a lathe would prove expensive. On the other hand, a part that crries a tolerance of 0.015 in. probsbly is more appropriately machined on a lathe or milling machine rathr than prenmon ground. The established dlaa of tolerance include run-
uwc 006q
l"l.s
2a
tool.
numne
PA
1.I14M
Rar,row.ntstlwo w~lthuJ C'• mr
with a built-up edWe. This action contributes to ort
ning or slking clearance (RC), location clurance (LC). locaton ransition (I) location interfereace
Wtool Wefe rqiring optinum tool omuetry. prop cttin speed, ad proper cutting fluid to nmne matriaM are ai. Iritt beet aw ndum removed by a comblnatloea of shear and freatue,
(LN), and force or shrink (FN). In RC, one part move insde the other. LM LT, and LN am used mainly for th eassembly of sationary pars. FN yieds comat bore prsum. The relationship 17-S
l4
I-U
AMCP 706.202 TABLE 17-2.,' UNIT HORSEPOWER VALUES FOR REPRESENTATIVE METLS
MATERIAL
£
HARDNESS
TURNING
DRiLLING
MILLING
MAG1••SIUM ALLOYS
40-90 Br
0&
0.2
0.2
ALUMINUM ALLOYS
30-150 B,
0.3
0.2
0.4
COPPER ALLOYS
20-80 RB 80-100 RB
0.8 1.2
0.6 1.0
0.8 1.2
TITANIUM
250-375 Br
1.0
1.0
1.2
PH STEELS
170-450 B,
1.5
1.4
1.7
CARBON STEELS
35-40 RC
1.6
1.4
1.5
ALLOY STEELS
40-50 RC
2.0
1.7
2.0
TOOL STEELS
5G-55 Rc
2.2
2.0
2.2
TUNGSTEN
321 Br
3.5
3.3
3.6
TABlE 17-3. "-RIPESENTATIVE SURFACE F1 IISHES OBTAINED IN MACHINING OPERATIONS
OPERATION
RMS FINISH,
BURN ISHING
2-4
LAPPING
2-8
_
_duced
-_
HONING
I
2-10
POLISHING
2-10
REAMING
8-50 5-150
GRINDING BROACHING DRILLING
.
15-60 _. 75-200
between the limits and allowances for holes and shafts for a Class 2 lit arc illustrated in Table 174. Among the more frequently encountered elements of design are cams and gears, keyways, splines, and serrations. Each gear type performs a specific role in power transmiuion. Gears may be cut on a milling machine or by hobbing, but more frequently are proby a shaper-cutter working on a preformed, forged gear blank. The types of gears employed are discussed in detail in Chapter 4. Splines, serrations, and keys are devices for attaching itcms such as gears, cams, pulleys, and torque bars to the power shaft in such a manner that there will be no interfacial slip caused by the torque imposed. Splines and serrations often are cut with a spline roller, keyways with &mill or shaper. There are many kinds of keys, such as feathered, gibhead, plain flat, plain square, round, saddle, tangential, and Woodruff.
MILLING
20-300
17-4
JOINING
TURNING
20-300
17.4.1
GENERAL
SHAPING
20-300
Joining operations include welding; brazing; soldering; mechanical fastening, induding rivets, bolts. nuts, washers and screws; adhesive bonding; swaging; and cable splicing.
SAWING 17-6
UNIT HORSEPOWER/IN./MIN F"OR:
250-!000
AMCP 1` 6-202 TABLE 17-4. VALUES TO BE ADDED TO OR SUB1TRACTED FROM BASE DIMENSION FOR HOLES AND SHAFTS TO CALCULATE TOLERANCE _____________________
CLASS AND FIT (Inthmuandths of an inch) LN FN RC LC ILT -
-0.9 -0.0
-1.4 -0.0
-2.2 -0.0
-1.4 4~.0
SHAFT
-.0.5 -1.1
0.0 -0.9
+0.5 -0.6
+2.5 -1.6 -
-0.5
-0.8 23 +2.0+.8
-0.0
+2. t2.28
change. Many parts fail in service because ol strris concentrations, whic~t tend to cause failure b) 1istigue even whed thc regions of stress concentra Jon are small - and are almost always traccable to impro-1.4 Per design or fabrication. The effect of joint d_.,:ign on -0.0 I stress concentration is illustrated in Fig. 17-2~. Joint design must take into consideration tti prior +3. 'istoay of the parts being joined and the pý.rticular *3.01 characteristics of the metal or metals. Thus highly -
-
0.2
2.5
-. 5
-1.6
3.9
-3.
WELDING, BRAZING, AND SOLDERiNG 17."1 Weldig Welding 4 the proem in which pieces arc joined or fused together, with or withouat a failler material, so that a cohesive bond is formed. There are two priMaly types of welding: fusion welding, in which molten metal is formed between the pieces to be joined; and forge welding, in which presure is applied to cause the plasticized surfaces to diffuse. A more recently devoloped fusion welding technique is electron beam (EB) welding. In this process. astream of electrons emitied by a hot cathode is focused to a fine beam by an electrostatic or magne17-4.
)
t..~~
-
Ia.J
the parint metal. Any abrupt change in the surfuce contour causes a stress concentration at the j~oint of
-
HOLE
ALLOWANCE
7(0o-l00. MIL-HDB9K-S, -693. -694 -697. -72 4.MILSTD-20. and ANSI Y 32.3-1969. Wherever possible, vWlde joints should I.: madc with smooth-flowing lines that blend graduwily with
-
cold-worked materials have lockoW-up strec
~es that
are relieved by the heat of welding, and t;is may result in distortion or a decrease in strengkii. The rapid heating and cooling of the weld can produce thermul stresses from expansion and contraction t.'it are quite large. Heat cracks are more apt to occur in weld metal than in the basn metal because the weld metal cools last and essentially is cast metal sith columnar grains. Corrorion properties also wiay be affected. For examnlc. although Type 304 stainless itteel is corrosion-resistant, chromium carbid.-s form in the grain boundaries in welded areas, lea%iii~g the rest of the grain unprotected. Compensation for such alloy changes may be provided by the filter ma,terial.
STRESS
CONCENTRATION~
-te -r no*iile. and
weldsi of high depth-width ratio, with corisequent lack of distortion, can be obtained. The operation must be conducted in a vacuum, and consequently is expensive and time-consuming. The proess isused mainly where high heat input, precise placing of the beam. and cleanliness of vacuum weldirig can be exploittet. The prcesa is useful on reactive, vacuum-melted materials such as titanium, zirconium, hafnium, and beryllium. It gives the same control of impurities a6 in the original material. This technique has been successful in welding titanium forgings; for the main rotor hub. Ref. 2 provides a more detailed discussion of welding practices. In addition, MIL-STD-22 dismacsse welded joint designs for manual and semiautomatic arc and gas welding processes, although it does not apply as a standard to aeronautical equipment. Additional information is found in AMCP
O
SmoOnh FO
~
.
-
Figure 17-2. Weld Coatove arx! Stress Conce ~ration 17-7
Joints shuldb located tothat the stire we'd
groove is vitcible to the welder and no obsUUtrcions umpair the acgeuuibility for welding. On all joints welde from both sides, the root of the first weld should be ground to sound nmea before welding the second side. Commonly used welding symbols ame givwn in Fig. 17-3. Representative butt joints are shown in Fig. 17-4. representative cirner joints in Fig. 17-5. and representative tee joints in Fig. 17-6. 17he welding procedure to be followed in the tabri' cation of a peat will be determined by the design and will be defined by the contractor as a proem specification. Tho cont and effectiveness of a welded construction will be determineid by the prescribed procedure. Pror to welding. the welding procedure and the welding operator must be qualified in ac-
17-42.
1auhms1
The Mainis of two pubt by brazing requires the ue Of nonferrous fill rod. tulna, or powder. Tb. pisces to be joined and the filler material are brought to a
temperature that is below the melting point of the materials to be joined but above the malting point of the brazing material. The brazing material wets the sufcso h icst e onI n houhcpl Wry action then draw the mlts Into Ohe sae. between the pearts. The optimium strength occurs when adhesion is pFset between the molecules of the brazing materialk and the molecules of the bm materials. Under these conditions, some alloyiarg takes p~ace.
cordance with MIL-STD-248. Qualification as-
surance requirements for welds are discussed in
Chapter 6. AMCP 706-2013.
B-14 f\
A 0O'TO 1/8"
(A)SINGLE-V BUTT JOINT, WELDED BOTH SIDES
B-41 T
soo8-46 L;tO M
ItC
(B) SINGLE-V BUTT JOINT, YIELDED ON BACKING
A, ELfhENTS OFAWELDING SYMBOL
FLANGEB-31
cw*.
JA Fl
IL
CROOVE(C)
A-W
91 ia BASIC ARCANDGAS WELD SYIMOLS FI
'IB-11 f
XK X( W 1 1ECBASIC RESIS IANCEWELD SMVOtS
Aý0._. 0 IEL SUPLEUTAR SYMOLSNOTE: (D'____________ELD_____L
17-8
Figure 17-3.
Welding Syinbula
TO 1/16"'
DOUBLE-V BUTT JOINT, WELDED BOTH SIDES JOINT
BALI,
uA.
'-""
NUMBER B-14
B-101
.ING
WELDING
dg 60
(m) in. 1/8
POSITIONS ALL
60 45
1/8 1/4
ALL ALL
X min, OP!E)N
FLAT, VERT., OVER 1/2 20 8-41 8-46 h 12 1/2 F L.AT ALL 1/8 8f-31 j60 8-10 MAXIMUM PLATE THICKNESS - 1/8 in. REINFORCEMENT OF GROOVE WVELDS SHALL BE 1 /32 TO 1/0 In. AS WELDED.
Fluvre 17-4. b~presematalu
St Joleb
AkMP 7
M2O2
T-I2 (SEE NOTES 1, 2. AND 3) .
1/8"
TTO
(SEE NOTES 1AND 2
'
.
(A) OPEN SQUARE CORNER JOINT, WELDED ONE SIDE C-2 (SEE NOTES AND 3)"
6 -
-TI?
S,T12 main
(3) OPEN SQUARE CORNER JOINT, WELDED BOTH SIDES
C-20
(SEE NOTES I AND 2) T
0" TO 1/8"j-1/8" rm
1/81-, mi
"i SF=••
'
o"
(C) OUTSIDE SINGLE-FILLET-WELDED 1'nnaRn JOII
r
(A)SINGLE-BEVELED TEE JOINT, WELDED, ONE SIDE, FILLET REINFORCED T-14 (SEE NOTES 1 AND 3)
Si2 --- %~1
(B) SINGLE-BEVELED TEE JOINT WELDED BOTH SIDES, FILLET REINFORCED T-32 ISEE NOTES I AND 3)
N
0" TO 1/16" 1/8"rmin
"mmA20min
7
C-21
S (SEE NOTES 1AND 3) 0" mEin Ey (D) DOUBLE-FILLET-WELDED CORNER JOINT NOTES: 1. REINFORCEMENT OF WELD SHALL BE 1/32 TO 1/8 in. AS WELDEDJ. 2. JOINT SIIALL NOT BE USED WHEN ROOT OF WCLD ISSUBJECT TO BENDING TENSION. 3. SIZE OF FILLET "S"SHALL BE AS GOVERNED BY DESIGN REQUIREMENTS. Fkgv 17-5. Represetatw Corser Jois
-
I -
)
BrrAing has the advantage that it can be used to join dlisimilar metals. The rivting point of brazing materials is above bOOF. Brazing materia include silver, copper and aluminum alloys; nickel-chrome; and silver-mangmane. Brazing techniques include torch, reistanw, induction, furnace, and dip brazing. Avnzity of fluxes is•umd. rits and tolerances arm of particular importance in brazing. Lap joints are necessary whenree strcngth is a consideration. The only allowable loading of
(C) DOUBLE-BEVELED TEE JOINT, FILET REINFORCED NOTES: REINFORCEMENT OF GROOVE WELDS SHALL BE 1/32 TO 1/8 in. AS WELDED. 2. WELD JOINT IS SHALL NOT TO BE BENDING USED WHEN ROOT OF TENSION. SUBJECT 3. SIZE OF FILLET "S"SHALL BE AS GOVERNED B. BY SIGN DESIGN REQUIREM REQUIREMENTS.ENTS. F11ure 17-6. Rip
ousilative Tee Jolsat
braned joints is in shear. An overlap of three times the thickness of the thinnest member gives the greatest elracioncy. Butt joints provide a smooth joint of minimum thicknew, but gre more difficult to fit. Scarf joints maintain the smooth contour of the butt joint. and at the same tim provide the large area of the lap joint. Joint clearance is the distance between the surfaces of the joint into which the brazing material must flow. For any given combination of base ond 17.9
filler metals, there is a best joint clearance. Values below the minimum clearance are weak because the alloy does not flow into the joint. Clearances beyond the maximum also result in joints with lower stmeigth. The normal range of clearance is 0.002 to 0.010 in. Brazing materials and processes, with approved specifications, are listed in MIL-B-7883.
rfastening,
rivets; diagonal pitch, the sracing between the Deaums rivet centers of adjacent rows; and margin, the specinl between the edg ofta pant and the vaitwline of the nearest row of rivets. For structural and machine member joints the pitch should be such that the tensile strength of the plate in the distance between rivets in the outer row is equal to the shear strength of the r~epeting siections of rivets. The rivet diamete dis uulecte so that d 17A42.3 Sederlag LK/1rto d - 1.4%476he~re iis the thicknessi of the. The principles and techniques for soldering are plate. Good dailg practice calks for plac4n tke much the same as for brazing, except that soldering center of the first row of rivets a minimum of oe and alloytk are primarily tin-lead base alloys that melt ono-hinl( rivet diametes away from the edge of the below 8000 F. Soldering is restricted primarily to plate. For multiple rows, the transvers pitch is 1.75 d joining sheet-metal surfaces and wiring for electrical (see Fig. 17.7). It is also required that:~ connections. Copper and steel surfaces, and surfaces 1. Bearing stress of rivet and plate be uniformly coated with copper, tin, zinc, or other compatible distributed over the projected area of the rivet materials. may be soldered. Aluminum and stainless 2. Tensile strus be uniform in the metal between rivets stee are examples of metals which are difficult to solder. Cleanliness and proper selection of tht solder 3. Shearing stress be uniform across the rivet and the flux in relation to the surface. to be joined are 4. Accumulation of stresse on rivet be mini-. r-ized. of utmost importance. No porosity can be tolerated in soldered joints. Electrical continuity and complete Rivet holes usually tie made .iveral thousandths sealina auainst fluids arm normal reiuirements. of an inch larger than the nominil diameter of the rivet. On comtptession the rivet e:parvdbi to film thic hole while also forming th drm head Puchn 17.43 MEHANCAL ASFEINGoperations may cause d~sgratao in the strength of Bvcause of the convenience or mechanical the plate surrounding. the hole. An annealing operation may be desirable to restmr strength in these parts t1a need nt-var be disassmble often are joined mechanically, with resultant poov j.istriarea. Structural arld machine menabe rivets usually bution of stresses and incieased weight. Such parts are made of wrought iron cr soft steel; but copper, well may be joined more effectively by welding, aluminum alloy, Manel, and 1rkndoe rivets way bb brazing. or adhesive bonding There asho is the quettion of which of the hundreds of types of mechanical fasteners to usc. The result often is overdesian. with Poor strmigth balance between the fasteners adtepurts joined. Mort-. over, there is little in the literature to guide the designer in making broad choices in mechonical fastcning. However, there are more than 10 Military Sp'oýWPcations on specific fasteners., and MIL-F;9700 ar~d MS-178S5 contain general specifications and U.andares for screw threaded fasiriers. Follow' ing is a discussion of various methods of rtiech'qnical fastening.
JJ
17.43.1 Rivetls For &lIaircraft applications, p~articularly whare dissimilar metals amc involved. rivets sAnil be Wewithi p~iiner. Ohme an adhesive is placed between the parts heing riveted in order to dampen vimlations and to minimiu failuire due to fatigue. Important items of rivet joint design are pitch, the spacing between rivet centers; back or transverse pitch, the spicing between centerlines of rowts of 17.10
X=FITCN DISTANICE
Figuei 17-7.
kusvet Spacing
.
I
AMCP 706-202
required where weight or resistance to corrosion are important. Although there are no gerer~i Military Sp~cirications on rivetinig, there are miom thui 50 Military Specifiations covering particular types of rivets. A general tiveatment for riveted joints it. steel. alurrinum, magnesium. and titanium isgiven in Chapter 8, MIL4IHDBK-S. m Waserssurface K%Mis, 17-4U Bolts. ~andnutsote arcwd wahen hejinswl Dolt otenareusedwhe an nus th jont ill not Ne permanx..t and disassembly can be anticipated. Maximum strength of bolted joints can be attained only when the grip longth of the bolt is at least equal to the thickness of the pairts being joined. No threads are to be in bearitig in the holes through the parts. Washers are added as required to permit tightening the bolt sufficiently to develop the load-carrying capability of the joint, All bolted joints must be locked or safetied, and bolts in critical locations, such as in control finkages, shill have two separate locking provisions. Setf-
locking provisions include nylon inserts; in the bolt or not I nekmna and1
-.-
) 16.If failure of ~
*
mnretviner&a
dg',piin elatioe
Chanti'r
helical-spring-lock and tooth-lock washers. Helicd-1 spring-lock washers have a du.i function. Va.'t, ' hcy compensate for looseness that may be developed during the use of a bolt or screw fastener, preventing losn oF tension between component parts of the assembly. Secondly, they act as haidened thrust bearings to facilitate assembly and disassembly of bolted fasteners by decreasing the frictional rcsistancc between the bolted surface and the bearing of the head or nut. Tooth-lock washers bite into the bearing surfaces and increase frictional resistance to motion. These washers are made of carbon steel, corrosion-resistart steel, aluminum-zinc alloy, phosphor bronze, and K Monel of varicis series. 174.3.3
Screws
Screws, particularly machine screws, may be confused with bolts, Actually, the equations for strength and the precautions on thread bearing are the same for both. ANSA B 18.6.3- 1962 covers both s'otted and recessed-head machine screws. Threads on machine
screws may be either unified coarse (UNC) or fine thri~en
(IINFI I4e.,ie hg'eht
nfe-mninto-eunkcrre i
a threaded assembly should occur, it is preferable for the bolt to break rather than for eithtr the external or the internal thread to strip. Thus, the leni'h of the inatng threads should be sufficient to caMr the. 14l1 load ncuiiqsar to break the bolt without stripping. The critical areas of mating threads are: 1. Effective crosssectional or tensile stress area of the external thread 2. Shear are of thiz external tlarcad (dependent upon the minor dmiaeter of tha tapptd hole)
the distance, parallel to the axis, fromn the bearing surface at the diameter of the screw t3 the largest diameter of the be,,ring surface. For scrows less than 2 in. long, the threads will run to within two turns of' the head. For longer screws, the minimum thread length is 1-3/4 in. 'The body is the unthreaded cylinder portion of the shisnk. The designation of a screw thread consists of nominal size (inches or number). number of threads per inch, letters of the thread series, and class of tolerance. On drawings. the designation may be followed by the pitch diameter toler-
upoi the major diameter of ihe external thread). When bolts are used to jo~in distimilat mectals, or to
LK- following the class designation. The formulas for tokcrancee and allowan~ces for the several series and
join materials dissimilar" frok'u the bolt niateriai, differenme in the coefficients of thermal expansion and in the temperature extremes specifiod for the hel;copter must be considered in caiculating the meximum stresses. Provisions alao mnust bernade for corrosion protection in such installatvons. Box w~rench clearances are given in Ref. 3. For wrench access, bolt centtio should be placed at a minimum distancc from obstruction of two times the wrench clearance. Plain washers are defined by ANSI B27.2-1965. These washers arm available in narrow, reguler, kind wide series, with ;roportieris designed to distribute loads over lare areaa of lcwcr-stiength mattrials. Plain wassers ame macie of ferrous, nonferrous, plastic, or other nu2terials. ANSI B27.1-1965 deines
classes of threads are given by ANSI W!.1-1960. The heads of machiAne screws may be recessed. hex, slottcd. round, countersunk, pan, cheese, or mushroom. Tapping and metallic drive screw types inc'ude rouind head, flat head, flat and oval trim head, uiidcrcut, N'liste:, Struess, pan, and hex head. Sh~ct-mrtal screws are defined in ANSi B 18.6.41966. Some of' these, when turned into a hole of the proper size, form a thread by displacing the sheet metai, while others form a thread by cutting action. There arc 12 types, each having preferred applications with shieet metal, plywood, nonferrous castings, plastics, etc. Sct screw~s arc used for preventing a pulley, gear, or other part from turning relative to a shaft. Generally sp.caking, 1/4-in.-diaineter screws will hold againbt a 17-11I
forme of 100 Ib; 3/8-in., 250 Ib; 1/2-in.. 500 lb; and Iin., 2500 lb. Scdf-tappitj screw threed insaui am~ hard bush~riv with internal and externa! threads. They often are used in rnmnicrorus castings. Hlcai%~ diamond-shared coils of stainless awlr with phftphor browce inserts oft~en are. used to rcziii old, thre-A.d holes. There arc many M~iizy Specifications *cair4 with psuticular typi- of scr-rva, including MIL-STD9 and M11L-S-7742. Additional data wkill be found in Chapter 8. MIL-HI)OK-5. pit)s, such Fas cottar pinf ae used to I==ur nuts upon bolts, nr urponi oftv pint and fas,=s. Cotter pins may bec obtained for bole ci=e of 3/64 to 3/4 in. The eotended-proxig type is sm-urcd in place most "eaiy.CNvis pins (ANSI HJ5.20-1958) fireq.&ently arc used -, lock bolls and may I,,- smcured with a cotter pin, Dowel pins (ANSI B5.20 t958) arc used either to parts in a fixed position or to prtzerve alignment. They normally arc wubjctl to shtearing stizin oaly at the junction of the two parts being held, and two usually ere sufficient. For purts that froque'ntly are disiusaexnbed, tht taper dowei is pteferred. This type ýklso is preferred foT joints of close tolerance.
Srctiain
17-4.4 ADHEWSIVE BONDING
-
-S.,..
~
~
~
(B) SHEAR
-J
--
_
STRUCTURIAL
In ei der to realize the maximum benetts of udhesive bonding, &structure most be dwtigned ini~ally with this method of joining in~ mind. The firmt requisite is an understanding of the basic loading con-
I
II
*
ný I JR tn 3 116 in are satkfi~a tni-to dm.v.tem
pin
hil ~ f&actol- in mlost Cases.tehl When soft parts are to be joined. h oeshudf 0.001 in. smaller than the pin. For Ilocking fit, astoiigitudinally grooved pin irs preferred.
4Aabout
(A) TENSION
()CEVG ()CEVG
d--.
hcpae1 structural applications of adliesives were basad on tension, shear, cleavage, and peel (as illuth a78)heiese critmeti pcurely sti~ual deiomnt oFig sta,, rath stviamtednt ofg17a)dhesie critmeri prelystulted applications bas&.d on design lozdtb. Today. adhesives not only must meet these requirements but else must include resistance to environme~ntal conditions cx~rm'rircde in current adhcsive specificat~ins for (D) PEL!.. resistance to moistuk , tactration under load (enviratimetital cyclic cerep) and undew dynamnic load4 it14 yk fJAda oo t"Jlf midJl4 yao oia Jlw 1(fracture mechanics and lr-igur). These include temir siom,, shcar, cl,ýavagc, a.. ' A*, as illustrated in Fig. causc failuore in the cienvage or p.-d siwation (Fit&. 17-8. 174(C) and 174(1))) may Wz desaribed as tensileAn icecal atnictur..! joint is co.c in wh.%h the load istresses, -dwy sic cocantrated heavily in~ local regions aisistibutcJ CS 611iformly as is possible over the entl of the j-ziits: thus, the Load capability of Jhe joint is bonded area. This condition most ni; Prly is Rpunrelat" to t!-c total bonded area. Thertfore peel pioached when fin: basic stess is tension or shetar and deavagc. situaticns 0bauld be avoidvd. (Figs. 17-6(A'. akid 1?-B(D)). While the sirescm that 17-12
PACP 70)0 The designer must examine carefully even thft tensaion and shear joints to minimize any eccentricities or deflections that would cause an unfavorable redistribution of stresme in the joint. For example, a simple lap shear joint under load, if rot restrained, will inaergo a doflection that causes high concentratiop., of stress at Points I and 2 as iliustrated in Fig. 17-9. The designer should keep ii. mind that joints wish uniformly distributed shear o: tension jo#,1ts are a goal, and then should apply coa-mon sense to achieve this goal. Such a common-sense approach to
Sdesign
a joint design problem is iliustiated in Fig. 17-10. The spar of the rotor blade might be consideied to be a rigid member. Aif loads acting upon the aft section create a bending moment that tends to pry, or peel, the lightweight skins away from the inside of 'he spar at Point A. Therefore, a "keeper" channel that gicatly rigidizes ths portion of the structurc is iticorporated, permitting the joint between the. skin and spar to ract the bending moment through a shear couple. Another typical rotor blade design is shown in Fig. 17-11. Again, there is a tendency for the skin to peel from the spar. In this case, the skin deflectinns that
(A) UNLOADED 2 j
4
-
"the
~(S) K
I'ý.
5
LOADED LOADEbe
-V
FIre 17-9. LUp Suhr. ,oeit Defl•caic Umnde Led
SP
.,
"
"
would aggravatc, the peel situation are prevented by the incorporation of a simple angik (A). The design of adhesive-bonded joirts aemands the proper mating of parts with regard to dimension and toleranc. It is imperative that aohesive-joined components fit each other without relying upon the adhesive to hold them in the proper shape. Where overall external dic.cnsions are critical, it is nocessary to account for the thickness of the cured bond lines in finai assembly. A reasonable thickness allowance per bond line is U.005 in., although this value should determined specifically in each case for the pa.,ticular adhesive and method cf manufacture. In the came of honcyconmb-sandwich structures, no allowance is made for adhesive between the honeycomb core and the face sheets. The cells of the honeycomb cut through the film of adhesive completely,
POIT
j
AFT SKI N
A.;
BALANCE BAR
KEEPER CHANNEL
Fiure 17-10. Ty~lcal nlotor Blade Design -
:
Alternate I
SPAR
Fi"ure 17-11.
Typical Rotor Blade Design - Allernane 2 17-13
AMWP 70&202
DETAIL A INSERT
FACE SKIN
ADHESIVE Figure 17-2. Houcyconk SandwlIc S&Yecturt
I
and actually contact the face skin. (Fig. 17-12, Detail A). The bond of f~ace-to-core is achieved by the fillet of adhesive between the face and the cell wall, wh~ch is another example of a &hearjoint. In Fig. 17-12, the insert might be an extrusion or any other form that results in a metal-to-metal joint, as opposed to a faceto-core joint. Thius, if the desired overall diwrtnsion is ~1.00 in. and.the face skins are each 0.020 in. thick, the(A
.
Ijllwjis
V
I
-..
aL
-
9%.
.
^
n
E!A t~
1166
.. L c
I
ISR
A
IGUamwing &ul Fu
two 0.020-in, skins and two 0.005-in, bond lines. The core, on the other hand, wiould have to be 0.96 in. thick, allowing only for thc skins. An exception to this rule is the case of the adheiive that incorporates a carrier, or scrim, which isconipote~d of woven or randcenly oricrnted fibers. In siuch a acm.th,; thickness of the cirnicr sheuld be allowed for bittween coic and(B faces.(B Fig. 17-12 actually is oversimplified in that the abrupt change in section be~twccn the honeycomb sandwich and tbr t~gid insert or adjoining member should, in moat "WA.c be mnade more gradual eCthW by shaping the jiolid member as shown ir. Fig. 1713(A), or by the addition of couublers as in Fig. 1713(B). In either cane, the ru~cs for adhesive thickness ahowancL: still apply. If a construction like thaz shown in Fig. 17-13(A) .s employed, a feaming type of adhes.ivr may be used on the portioit of the core that fits against surface A oý the closure. Such an adhesive cat. be preted intcv the core cells prior to asaea1~ily and thereby permit fitting of th, ,arts. Another examzple ot a des~igr technique that a 3urts a fit of part and, thus, a uniform, uristrained bond linec is that employed for the balance bar shown in Fig. 17-10. The bar is bondN-4 into the nose r-idius of the C spar. Because the ap~ar is *ary stiff in this highly curved portion, while thý. balance bar is solid, tnec required bonding presuire it tchieved best by applying %force apiinss the back -ifL~c bar. It will N; noted that the nose radius of the Lr ipt 1-irger than the 17-14
SURFACE "Ate
-
DOUBLER
IW Fu
obest Howeycob Structure
71.Aal.o
inside nose radius of the C spar. This asires that the bar will never bottom out and, therefore, pmessure always will be applied at the side of the bar. This technique is not restricted to rotor bit".s but can be employed in any situation where the bonding pressure is provided by a component of force that is applied in a direction othat than normal to the bond line. While it was stated previously that parts usually should be formed to fit viah other without the aid of the adhesive bond, the degree to which this rule should b,- imposed is depentlent upon the stiffness of the members beiog joined and nature of the uammnby. General'y, sheet-m"t' parts should not be so closely tolerancod in thoa. free state that thus is no allowance for elmakance w~hen assemzbling thiem. There also must be space for 0.0 10 in. or more of uncured adhesive between the cowmponensl during
**1 *
-
M
ammisil. This can be achieved by an over-bead or wader-bead allowance, but this allowance neve should be so great as to prevent hand pressare fronm bringing the pert imto the final form that isdesired in the flnslndauubly. The rm~idual strains caused by such a =odto r neglgble,
When dissimilar metis are bonded, merious distortions can occur due to differential exparisions, because the bonds are being cured at taunpetature ranging from 226* to 3500 F. Such distortions of the structurre can be minimized if IhM member with the lowest coefficient of expansion is stretcbd or
Donded structures usually %re assembled in ~tooting. Thbis precldes a visWg examination of the reative locutions of components during the curing dos. Dmiase the adhesive becomes fluid at some point during the cure and thermal expensions are taking place, some shifting of ono part relative to
otherwise strained by the application -)f external force while undergoing cure. This often Lueau that some extra length must bc proviied to permilt gripping the member or pinning it to) the bond"- fixture. This extra material can be removed after the banding is compkted.
another usualy must be aniticipated. Thbus, adjoining parts must be dimenioned and toleancemd to accommodate such movement. Fig. 17.14 illustrates such a -mme The bottom of the figure shows apap where two members butt together: This gap must be programmed in the design and usually should be from 0.0W) to 0.00 in., depending upon the size of the structure. In the example shown at the top of Fig. 1714, insufficient allowance was provided and the resulting joint either will have avoid or. atbest. w.l be suhject to -ee. Usually, the small gap ilp the proparty designed Joint will fill with adhesive squeeze-out or masy be fiAle Lawe with a fulnnig-sealunge compound-will Currently, all structural bonds require pressure during their wring This pressure must be reacted it, some way. Eithe one of the members being joined must be stiff enough so that it will not deflect signirEcandly under the pressure required to boind the other munbers to it, or provisions must be made for tooling tbat will provide the reaction force. W; ;n such tooling is required, the detail detign must provide space for the toolinig and a means of removing it after the cure cycle is completed.
It often is impractical to complete an entire as, menbly ina singlie bonding operation. This create& Uj necessity for secondary bonding, in which a part of the assembly is reheated to bonding temperature while additional parts arc bonded to it. The most commonly used adhesives are quite weak at the temnperature at which they were cured and, thus, there is a risk of destroying the or~jinal bond during a secondary bonding operation if proper prt~autions are not t&kcn. One such precaition is the use o a lower-temperatur@ curing adh,,4ive for the secondary bord. Another, more reliable technique isto respply pressure to any oi henprimarily bondr..l joints tihki be subjected to the heat of seccndary bonding. Provisions kmust be made in the design to assure thaý resoplicstion of pressute is possiblet in such case. Finally, the desigaer always shouldf consider pro,viding, for physical tesking. sorn* kind of 'ntension of the basic straurc'~ that can be removed after bonding. Nondestructive testing techniques are being improved constantly, but there is no substitute fordrtructive teits of joint.,that ere built alont! with, and duplicate, the actual structure. CBI. ri,
CAW
Wire rope and cable may be used in helicopter control mechanisms, although push rods are preferred. Aircraft cable made of high-carbon steel wire, electrolytically galvanized and drawn tco size, has the highest strength and greatest resistance to fatigue of any cable. Representative wire rope fittings ume illustrated in Fig. 17-15. For aircraft, the more commonly used types range in size from 1/16 to 5/8 in. SwAged fittings on wire rope: have a strcengt rating equiv~tlcnt to the strength of tht wir.t rope. These fittings are applied to the end or the body of wire rope by tht application of high pressure, causing the steel
I
~IMPROPER
DESIGN
0.020 in. min INCLUDING TOLERANCES ACCUMULATED I Fkmur 17.14. Balance &ar Dealn
to flow around the wires and strands of the wire rope to form a union as strona us the. rope itself The ~necessry high pressure and flow are axomplished by means of special dies. Machincs for this purpose are described in MIL-S-6180 and MIL-S-8035. 17-15
"
0
SINGJLE SHANK LAL&L
EYE
NTRAP EYE
U41OLT CLIP :%amw 015n.
*
:.
-
90
DOUBLE SHANK BALL
EYE FORK
STRAP FORK
THML
Aircraft W*. Rope Fktthg
FgUre 27-16. CaW4 Sp0ldq
Anathtw eqctivc racthod of attaching fittings is by meais of p.tFZ3 zinc.. A special high giade of pure z~nc U*used tc* 's! ihc socket. (Babbitt and other adloy!, wifl not hold proporly.) When prc-pzrly prepared, thc oi .stm-noth approaches that of the wire
cable. Normally, thrcu lwr&m for ca~d 0=r0n are all that can be ,nana~d. Splici is illusuateJin Fig. 1716. Field-spliced conilu. never should 1be cxposei to moure than "0 of dwI w-cmal br'ahh-ig load for thte cable.
fo.: tcitpiiary eri.rU-bolt clipi mr'y be employed. W f :.nn a loop with a U-bolt olip, a ai~rope. tbimtlic 6houM he placed in the loop to prethat the saddle or base o; vetkin~king . ;t4 ewq 6ss~.iu! the clip boan aga~tst Ithe lon~gere i:livc, end of the rope, whi.2i the U-b4o'% bonr5 agfiiist the thortnr or dead eaid, Vhco end of t!hL: wix rope. should be seized pr.3perly. The strcngth of a dip fvuatenriig is zz-_ than 80% nf the etrenzb-' of kh cable. A seldorn-used meva '1cjoinian, cables and makii. endloops is cable 5jZ~*The method may be ewI in the field for repair whem equipment or fittings wr~ unavailable. To join twc, ables, each of the free eiud strands is warkeA ovar And under a stranid In~ tie othm cable, workinj axmanst the lay. A total of four tuci.; for vA6~ atranA 1a adequate for mrst purpovsi. An sys splkwcAew be made as a loop of any size, or 4iht arnand a thimble, The 6ads aft tucked under alw ioian stiraud& ftainst tbe lay and around the
sheaves or drumi ýhuw nn too swafl in diam=~ arm the principal resuotu for dwcrioretion of wirr rupe. A kinkedi, bent, or flaftencd ureA cannot bie Woeasted in a controll cable. Thw rope or cabIk should bte kept weil 4ubricated, usingj '1~e l-Aticant 6qurplicd by the manufittfurev. Normally, lhý lubqant maust be Imtetd to 2WO'F it! oro'ý tv peaic4uo. Nylon o,. vinyl coatinip are dfecrivc for reoling in lubricapt and for protacting wi -ope from dirt =4 o cric~'n.
17. lý
17-5
HE~AT TREATMENT
17-31 (1ENERAL ltat tteatment is define~d as the application of firav-temperacute-cooling cycle relationsr'ip. in order te cause atomnic, molectd~ar, or crystalline tramsformuations in materials. These trainsfotmations ame selected to impart desirable naroporties fujr particular and-oses. Although such tirawformadomi occur in
plastics, commica. and other Wonmealic mateumuis to astala. The engineerins measinchude feru OWPU
Hea ts
i 51)s,5WWUMtt5iU. n insosapoied toall frms of moa n
"MO,madone patsbwa. dwft f~p,
plates anti sheasa. The heating cycle may be con ducted in various tYPes Of furnaces writh avauriety of aimosploese in Sqi mei such as moe wit or molten le0d or by some other mean. The holding sad MiOWn Cycles may involve vacuum, controlled limrt ztnosplowa4, oil bath quenching. air cooling, or wone quenching. ad fUatulities- eqwplomtm
pr wt,
Qss.cVtols, n
eauptancs "taWards for the heat ftreaing of nedal parts have been define in nWuniems Military Speciftcadons including MIL4I.6OS,
Cutn
oddt nw
rvosydww
rsA
374JU. SamRlr This proc.. is conducted at considsrably lower
temp~erature than those for 0%e other host treatment Proesss.
tresses imposed by forminag or machining
we relieved by allovwinS the diffusion of boundaries,
or the diffumion of hydrogen from a part whome
hydrogen einbrittlwenuM is &batad.
174.2U Teuipeflu This is the Promes Of reloeating a normalized or quench-hardened alloy to a temperature just below the transformation rang and Ume coolisg it at asuitabbe retw TemperWSis used to obtain desired propertie of strength and toughness
.6875, -7199. and
41200.
174M2
174.2 HEAT TREATM104T MWTALLURGY There ane many kinds of heat treatment, with tunetemperatur cycles deupgned to develop te ciyua~ling -90Wran itr.c... f _J-
Aging is a tempering promes in whicli certain alloys swe held at a constant. relatively low temperatumI lonig enough to permit the pmncipitation of par-
I)
Alg~n
tices and grain structure transfornations that dewlon the dwirmid urnets of starentb. tough-
~ urndnde woess, haron.., or formanuiy. In some alloys, agant naeriaj. Thes characteristics MI~y bg 11CcOniP11iihd at room tGIperaUre. toughness, hardness, machinability, freedom from The data in Table 17-5 are indicative of tIte residual stress or any other property noevess'y for the parn to perform its design function. Each "&Ia teflperatures at which heat treatment operations arm Conducted for Ceacil Of the C*DMW4in MMrA)A. and metal alloy bas its own head treatmnent character-
led. thaL are determuined by its cbesica constitudon.
Tie morn Commonly employed heat treat Opera. U~ms am annealing. GormaliaiQ&g amus relief, tempering. And 161". A ~
7JFROSALY 143FROSALY -,U conditions for annealing hardening, a;W tempering Inividual feruous alloys arc given in the applicabole Military SpocirrAation or Aerospace
Moatheat processes ame variation of sanwali;% A "tad anneal consists of heasting the metal to a
fiterattim. Air, vombusted Sake, protective atmoepheres, inert atmospheres. vacnum f useals and molten metal ame acceptable bcati*S media. Tanperatwea control must allow the entire lot of het-treated mtatial to acidlve its desired properties Corrosion contamination, and embnittlement of the metals being tmiaed sAd! am be permitted. Procomme such as induction beating, flaoe hardeniog. waburiong nitfridin austemnpeing, and murtompering awe rocognized processes but anc beyond the saope. of this diecuesson. The equipmaint for beat treating musat provide for adequate handWin and uniformi treatimet of the charge, as well as for prmclee control of the timectempecature-cooling program, to a..lmiv the desired results. The quenching proome involving oil, Water, and sir must be controlled to function with adequeai speed. 17-17
Imq umrtu thetallows the grains to recrystallizeins duk~d p1101n. This PVoc. is descibe inter-
donosably as soltionm heat umatinet. to indicate 2Wthe meta it betdW to atsuoisrature at which Johe alloyng elamelas an dissoved and placed imsolid solugon Thai the nistat Is olD in Omder to procipiate the desired etructure. 17-S.&2 Nweaigul T1his annealin step oftes is accomplished after forgiog or machining in ceder to restore uniform. grain structure. Tho meta Isheated to just above the wacldon tempoerau - in no cose long eniough or on aougim to camr tbo Savaw to gpow to any aputeatmo - OW dote. b. cooled in still air to - temperuttre. The effect is to restor uniform
TAMES 174. RKVMrnNTAJW FE? IDTEAl TEMPERATURESC'
METAL STEEL-041' TITANIUMI
4 A1--3Mo-1-V
COPPER-Be ALLOY no 175 ALUMINUM 2024 *
TEMPERATURE IN0OF FOR: SOLUTION H T NORMA LIZE ANNEALING FURNACE 04Y0L AiR COOL 1525-1575
*
oo
QJUE NCH
QUENCH
STRESS RELIEF AIR COOL
101$5
1600-1700
j900-975
1620-1700
i700
90J
900-913,
370-380
The foliowi4 limitations and coetroke anv imx'*
TEMPER
16007l,77) 1525-1600
MAGNESIUM AZ9C7-75-79_____
*
jHAR DEN
the tresima nrnm_-nm
40W425
ianozrsiou iWwater of 1000F for m', qbt alloy xWi IS60 to 212*F
1. r~nfoxatioij~enig teca ailbc qucichd to noitwls than 93% raatewaot 93% lowaz bvainite, as specifwed by deign docwgciaiatin. Mixed Cu~ctuvos an not acceptable. 2. Cooling ol &~ahee steoLs froml the mnneaW rw~e Mmus not cxcoed W0 des F/br doiwr to Ii AI of.
t
900-1100A
_____
far es~nauswso an4 (ns.
mnafl that
certain allys maý be osl-ucache or. air-quomdcaf. 17443I Ca~er AMeys Fuenaces having vazcuus or cowtolled amo-. apl\vxw fresquently ame used for~ ot'qw. Air atmoqýberc may be used whmtien lwekm of material due
Tr~orn~onardn~aea~ibecolsd to oxidation and scaing i& vas dacibmuW to th to o belwqtx~iikbat terpowu befre fi-ished paut. Bright batrdcsdqg requins a oceuollsed, 1bcr maximum ptidisuibkand cnoe inwdcpdta.a£ajSonephe in 4. lzrv-.mssik mximm iwuw n dxh . achamber furnuai. Moltqn cat bats_ APU aw be any zone of &r~rburizn& 3&Ul ano ecated 0A003 in. used for solutioni has treatment bwc&isw of the urns -- ----- -'-~
4
44,OAC
41
-,-------'
tquuaay us uVAMmw-iiyiis
wxiwwm
unew by
nuiclii~ig '~i~t~s-molten or Saw ntadvt s ion hea trmmsnoqt tagratuggs * 174NNFL~cOU ALOYSTinme-teampenture-coooling cycles mut he aequately 17-A NNFUOU AL YScontrolled "od the pats adsquauiy chamnd befor 114A.1 AbmJ'mWm Alloys charging. Cleaning masy involve ivpor dagsta, acdd Air. fiavidized beds, combxusatd gnus, protective atpikling, or- briabt dippin. Novtra uwi balbs may be cxasptraa, adni olten salt bafts arn acceptable employed for Age hardenin bust mwAi be rmeirod caitfully and neiztnliacd. qziemntizg is n wD atr. aiedi;%for tL, heat trcating of aluminum alloys, proCopper-beryliium mill prndu,4a and fou~i;%a~ x vied %haW no damage is done £o, the materia. Iranu mally are sumpbcd in a con'i~ton suMA-t ft pt'~iiiipIoWAsd in air chamber furnaces must be shown by tw tol be free froxn bigb-twmpurature oxidaion. Salt tation beat trixat4 bnt, so olution beat l1tattng ba Wiutt~ muss be of the proper type and grade; nitrate performed wnly wbei wa~iia or cold workLg has mn baits will atekack aisnuinum- magnasium alloys, for qwirod a softening trearauct. Ti-npta coo&%n cycle controls lly 17&4.43 TamA masi provide the desired puopnissf. Part-, must besAbwtiunttly (in LoK% lubnicanz arAd oth foreg Fumraes having a slightly oxidiing or inert atnvfter which coud harm the inntcial being heat naosphcre ane employed for boat treating of titanium ts*4&i. Quenchi*ng ormahAy is zonoducted by totW allys. Reducing or cedodrermic atmosphere such as -t
-
*'1 hydrogen or cracked sammonia AMl wenbe used. Hydrogus ambrttleamnt presents amajor problem beasuse hydrogen isabsorbed readily from baths and eSum Queuackiag Isin water. and quench delay times must be minimal, exempt that th, product sAll bL air &xole aftier striess-relieving operations. The treatmean of titanium at temperatures above I l000F uander oxidizing conditions may result in severe scaling and oxygen diffusion to form a hard, brittle surface layer. Titanium alloys are susceptibie to stress corroeson by halides at tsiperatun above 5500 F.
hardening. And, as a consequence of their high duotility, much work hardening occurs before rupture of the chip. Th6 restilts, ir. increased power consumption and tool wear. On the other hand, work hardening can be beow ricial. lacr.ase in tensile and yield strengfth can be obtained. as can resistance to bending and buckling. along with increased fatigue resistnc. Judiciously employed. the phenomenon can be used to provide lighter, stronger parts, thereby eliminating the need for subsequent, expensive heat treating equipment and processes. The designar musm be careful to select 1745 D=SGN ASINCTS 01t material that will work harden to the proper degree HKAT TREATING during fabrication in order to produce a part havis-4 the required hardness or rigidity without excess The beat treating processes to which a particular wetght. Knowledge of which gape and condition of part may be subjected in this course of its falrication material to select can produce significant savings. For and nassbly aue an integal pa:-9 of its doWi. The instmance, many materiala can1 be purchased with hbut treat requirements must be defined dlearly on the drawings and in the process specifications for the, varying degrees of cold working; aluminum, stainpart. Ihe hrAt treatment must be performed in thme less stel, and brass may be purchased fully annealed, 1/4 hardened, 1/2 hardened, 3/4 hardened, and fully proper soquence to achieve the required end-item hardened by cold working. properac. Because uj. end-isan properties serve to qualfy the ntrire procew the design must be succh 17-6.2 FORMING that the properties realistically can be achieved. Assihinse au "M mm %av4"45% m -. Of '-W, FrinUngD uptieiuli i-w ae-A* i tah.ý*S-j metal parts can be found in a number of military where savings can be obtained by judicious use of the documents, including MIL-HDBK-5. 4693, -M94 proces of work hardening. Ali operations that in-697, -691, and -723. Additional information will be corporate bending, stretching, or upset of metals, and found inm Rsf. 4 mad in Chapter 11. AMCP 7l06. 1 result in plastic deformation below the crystallization temperature, involve work hardening. The efmay be advantageous or disadvantageous, doIGfeag WORVH"MC 17-6 174 AR~fINGpending ORY upon the selection of the starting material and the rate and degree of deformation. For instance, 174L1 GV4E3AL stainless steel has high ductility but wrinkles easily 1 qWUpona Instal, sOme Or all WbeaaZse0miWes with compression. A strong, light muffler header having incrcmsed rigidity and increased resistance to inntip lIsvium zW~ft;1es. no"lu iifth iowF, h adaiiCow WUI fasis" mpe can be macic by nmiual "tigeomuh &rmfim if eeu hn~ mmamsel, th atms wll ~ .e~ ~of stainless steel and then designing tht blanking and fatic p~aooy U atom has no reu t nhi forming dies to pro .'ide the required degree of deformation in the proper places. In this cawe, an inner formation. W"'w this occurs below the recrystalstretch and an outcr compressive deformation would lizastio tempesture the metalisa said~ to bie "cold worked"; iLe., the grains have bo distbetrequired. pid, ad fagmeted As ~mmi~cfedthe i.Similar considerations also may be appropriate for dished and flanged parts such as wheels, pulleys, and in the crystals move to grain boundaries or otheimerfetios, werethe arestaled nd ,efairings. and in the shaping of bars and tubes, intcset icresingresstace t fithe platic(Is gral stiffeners, and large, stretch-fovnied shapes such formation. The metal then is work hardened. a olns Work hardening has some distinct disadvantages. 174.3 ROLLER BURNISHING It can cause cracking in sheaet-ntal forming, or it can require intermediate annanithii sups during shaping Roller burnishing is a method of improving finish opmi~vos. In machinng, the metal chip may besod dimensional accuracy. and results in work com sverlydeformed before breaking away from hardening a surface without the removal of metal. thme wo'k piece. Austenitirc steeb are difficult to The operation is employed primarily with internal amabine because of their high rate of work borns. Donv diameter can be increased by 0.002 to
"4
*locations *
-
17-19
.
AMW MMG 0.005 in., although this is not normally a primary objective. The operation firequently is designated for phosphor bronze and sintered bronze bushings. The depth of burnishing normally is limited to three times the diameter of the hole, but the insides of tubes 10 to 20 ft long have been roller burnished. Wall thickness is limited to no k& than 1/16 in., unless the wall is supported properly by a backing matcrial. Metals that work harden rapidly must be at a lower hard-, ness. Roller burnishing is a machine operation in which a set of steel rollers is caused by cam action to impact a surface at a rate of, perhaps, 2,000,000 blows per min. This produces a smooth surface, improves roundness or straightness, and increases surface hardness to a depth of 0.005 to 0.015 in. The surface may be finished to a tol-rance of *0.0301 in The kneading action tends to reduce the stresses imparted by prior operations, such as welding and machining, and also introduces a compression stress to the surface. Greatly improved fatigue and impact resistance can be given to parts in this manner. Roller burnishing has a limited amount of application to external surfaces. A special operation using a na.,--•,
,.llor •r
t
•f.
-.
ime ..
is
employed for rolling fillets. In this case, the objective is an increased resistance to fatigue. Relatively small forces are employed - 100 psi or less. A plain roller of oil-hardened tool steel at Rockwell C-62 to C-65 may be used to burnish a fillet of 1/32-in. radius in about 10 passes. The rolling and pressing causes a combined rolling and sliding action on the metal in the fillet to relieve the stresses and to work h'rden the material locally to higher hardness ard fatigue resistance. 17-6.A
SHOT-PEENING
Shot-peening is a process used on many helicopter components to increase fatigue strength. Cornpressive strcsses are induced in the exposed surface layers of mctallic objects by the impingement of a stream of shot, directed at the metal surface at high velocity an. - under controlled conditions. When the individual 1,articlcs of shot contact the nectal surface, thoy produce slight, rounded depressions in the surface, thus stretching it radially and causing plastic flow of surface material at the instant of contact. The layet of metal thus affected is 0 005 to 0.010 in. thick. The surface metal is in compression parallel to the surface, while the underlying metal is in tension. The compressive siiess may be several times greater than the tensile stress, and therefore offsets an imposed tensile stress such as is encountered in bending. The fatigue life of the parts in service is improved marked17.20
ly. The strest-concentration effects of notches, fillets, forging pits, surface defects, and decarburization are redu-ed greatly. Shot-paenin3 a change to ben*ficial compressive stresses the resiual tensile stresses that grinding usually imposes upon a metal surface. A higher residual stress approaching the full-yield strength, can be obtained by strain peening. This consists of peening the surfas, while it is being strained in tension. The surface tensile stresm that give rise to stress corrosion also can be overcome by the cornpreuion streams induced by shot-peoeing. The brittle failure of a ductile material due to stress corrosion has been associated with brass, stainless steel, aluminum, zinc, magnesium, and titanium. The cornpessivc stresses due to peening are stable in low-alloy steels to 5509F and in high-temperature steels to 8000 F. The effectiveness of peening in improving fatigue resistance is ilh'stratcO in Table 17-6. Shot size has been standardized by SAE J444, and the shot numbers range from S70 to S1320. The shot number is approximately the same as the di, meter of the individual pellets expressed in ten thousandths of an inch. Cast steel shot is the most widely used peening medium. It has a useful life many times that o,Ct ..--
-a ....
.
.
"d
wJ
o-
components of peening machines. Cast iron shot is used in peening operations requiring low initial cost. Where contamination with iron is not desirable - as in the peening of stainless steel, titanium, aluminum, an, magnesium - glass beads are employed.
17-7
TOOLING
17-7.1
GENERAL
Tooling for helicopter manufactare is the responsi..n.-., an-.. -u.*aenr*,,..r a., •-y .... a. -- -C ... • ., .. fined in the contract with regard to Governmentfurnished tools, is unique to the manufacturin2 facilities of the manufacturer. In any event, configuration control will be in accordance with MIL-STD-480. Tooling is a significant element of helicopter manufacturing cost, and the more stringent the manufacturing tolerances, the more costly the tooling. Consequently, it is imperative that the tolerances specified by the designor be kept in perspective. In modern production work, where mating parts art manufactured in different departments or by different contractors, some method is necessary for producing these parts so that they will fit correctly in the final assembly. Appropriate standards include MILSTD-100, ANSI 84.1-1967, and ANSI Y14.5-1966. As with airframe design, tooling design mast be performed in accordance with standard practices and
AMCP 706-202 TABLE 17-6. THE EFFECT' OF SHOT PEENING ON THE FATIGUE PROPERTIES OF SELECTED SAMPLES
STRENGTH GAIN BY PEENING,3
SAMPLE PLAIN ALUMINUM Z014
-
T6 ROUND BAR
23
PLA'je ALUM'INUM 2024
-
T4 ROUND BAR
34
PLAIN ALUMINUM 7079
-
T6 ROUND BAR
30
SPRING STEEL 5160 FLAT LEAF
51
PLAIN STEEL 1045 POLISHED
10
SINGLE GEAR TOOTH 4118 RC60
29
S-l1 STEEL
54
-
GROOVED
0.54 % CSTEEL -V NOTCHED 4340 SZTEEFL POiSE m r
UnD Bnm AP
procedure. Tool design ztandards, prepared by the manufacturer, should include design drafting practices, design and shop techniques, standardized tool specifications, types of mateaian, standards of material strength and dimensions, and tool production and qualification processes. For most projects, tooling will fall into two stages: nr
S) '~
fm
73
tale~Ine anti ,wnrd-~sinn tnnfine Wlwpn the
designl of a helicopter is released for masnufactume drawings and specifications for all of the parts and assemblies are used for tool planning and tool design. Tool enginees consider the number of units to be produced. the required rate of production. the equipment and resources of the plant, and contractual lirnitationst, if any, on tool costs. When all factors have been weighed, the, tooling plan will be defined. The prriduction planners then can break down the manufacturing processe for eaclh part into indvidual operations. The helicopter will have been desgedW by qngineers who are concerned priuiarily with ttw: propet functioning of each poirt, although they will have kept ill mind the factors of producibility and uconomny. The tool engineer, familiar widi the mianufacturing facilties and the tool stc'ckpilc rvan will consider me dsignof the part for easier Or trnore economical
150
manufacture. New material coipousitions or cot'figurationio, new manufacturing processes, and new tooling sequences are possible. F~or instance, a cast ing might be replaced by stamped sheet metal, requiring a forming die. Otten, a suitable tool cani eftec~t economy in mnanufacture by reducing the amount of material scrapped. When thic mrichine sanuence hps bam ettrMined. the individital operations are fisted. An operation consists of ail of the work that can be dcne at one setup, or Itation. Thee operations are planned in an or-der that will reduce the number of special too.ls to a minimum. Thus, it is better to design dies for multiple operation on a single press than to require individual operations on a number of punch presses. The same niultiple-use capability is desirable for jigs and fixtures. Once an operation is listed, the tool to be designed is determined from tht description of the given opera. lion, the machine to be nasployed, and a set Of dttailed drawings of the part to be made. This tool complete with assenbly drawings, subasscznblics, part details, and specifications - becomes an element of configuration contiol for the haelicopter. For both prototype construcgion and production, theat are three broad categories of tooling: shop 17-21
j
tooling, airframe tooling, and tast tooling. The latter
made and the material the ports arm iade from areAl
two categories may, in turn. require shop tooling. 17-7.2 SHOP TOOLING Cutting operations in the shop are performed, for the most part, with standard tools or with ex.. pandable tools that do not become part of contract inventory. On occasion, however, special bar tools may be required fo boring, reaming, mecessing, grooving. undercutting, and similar operations. Most of the specias tools for shop use are jigs and fixtures for holding the work piece and guiding the work performance. 17-73 AIRFRAME TOOLING Fabrication of helicopters requires the transfer of numerous dimensions from an engineering drawing or prototype to the item beinxg built. Where the number of items to be built is small, and the performane requirements not too great, the. structure can be produced economically by extracting dimensions directly from a blueprint. However, for the . - . -. -'.
cise tools have been developed for transferring the configurations of nondimensional. lofted, full-size drawings to the actual structure, even thoukh portions of that structure are manufacturn by many different manufacturers. The five primary types of tools tmployod ae: 1. Master tools, designed so that dimensions for the entire helicopter can be referenced to several master tools. Each tool is the three-dimensional representation of a key portion of the system. 2. Templates, a thin plate of metal or other suitPieo material that may be used as a guide or pattern. A template generally defines the profilr, contour, or layout of holm; the bend lines of a part, or an assembly layout of several partL 3. Optical tooling, a system of tools constructed with alignment telescopes at one end and targets at the other end. Witt the use of such line-of-sight tools Stogether with master gages, tooling bars, increment bars, optical micrometers, and a transit - the configuration of a system can be conolled from a single datum line in six degee of freedom. 4. Jigs and fitures, coordinated to the master took, position and hold the detail pars in their eltionship for drilling and fastening. Smaller or subassembly componentr are fabricated in fixtures that are coordinated to larger or main asmnbly Fixtures. The large components then are loaded into joining or final assmnbly fixtures to complete the assembly of the airfranwe. 5. Plastic Tooling. The number of parts to be 17-22
...
determining factors in the type of toolin to be used. Then are three common elements used in plastk. tooling: a. Mock-upls (1) wooden s (2) piAsto filled b. Molds: (1) high temperature epoxy (2) low temperature wood (3) matched (4) aluminum shell c. Trim tooh5. Special tools in tbAse classifications may be designed for inapectior., or for the application to tting of fabricated parts or system3 us!ng the test tools described subsequently. Basic decisions as to tcolinig design must be made concurrently with design of the airframe. Tooling drawings will be initiated as soon as a flow chart and production breakdown are possible; the basic tool philosophy and procedures must be established at .!
-
-
-W L
...
-
..
-..-.
1
ý .
A
- - -
.
.
allow maximum latitude for airframe deign changes while minimizing the need for redeign of tooling. Continuity of datum points and coordinate refer"ce lines and planes must be retained at all times. In most cases, the tooling proemses employed will involve all five types of tooling discuseed, but will rely more heavily upon one type for configuration control and quality assurance. The quality assurance program will be determined almost completeiy by the processes selcted since the smun system are eraployed for inspection as for production. 1 be introduced when Considerabie variat -numerical control loft... - employed. In thi cas the engineering data are transmitted in mathematical form. Automatically programmed tools are possible, with a computer working from, or producing numerical-control drawings. Provision must be made for special situations. For exampl, becaum of the small cross section and the extremely close tolerances required in tk: menufacture of rtor blade components, normal lofting practice may not be followed. These components may be fabricated in closely dimensione detail. Rib and other blade components may be lofted actual sie and provided with appropriately toleranced dimv sions for inspection purpose. When justified by the nur ýer of units and the number of subcontractors, the manufacturing and assembly processes may be built around the use of master tools. In this cam, quality assurance can be improved and simplified because mtew tools slint-
|
ans individual intespretatlom of enginsering valuta and permlt duplication of does toleranne. Further,
bration inputs in specified attitudes and types of sacpnion, fatigue of auemblies, end determination of
by having a coatrol matter as the dlinusiomal au-
rotor stability characteristim.
thomlty, duplicate master can be built for uie by subcontractors. PFonection also is provided agains the Ion of dinvmmlonal control d•old mowe tools be damaged.
Desig and conAruction of special test tooling must be initiated as soon u the final configuration and the qualification specification have been deqted. Verification of the test tools and test tool procedure must precede their application to the evalusLion and qualification of the helicop and its subas-
17-74
T'10
TOOLING
Most pbys•al, mechanical, thermal, and elecrica Ntuing will be mcomplishad with standai input and readout equipvient, and by un of expedable apps.raius stuh a &*=in gags and relstance strips. However, some of %,equalification requienmeis for
samblka and assemblies. . RMFllENCES
a blkMopter. and its ammblis or subemblies, wquire testing in a configuration or in an environmental or fatigue condition that is not attainable with stmdard tes equipment. Such tests, for example, may bacom wneesry to design and fabricate tst fixture, tNet stands, load input equipment, and readout equipment peculiar to the helicopter system that is to be manufactured. Such special testing might involve the twisting loads on the helicopter fuselage, vi-
I. Meaal Forming, ASM Handbook. Vol. 5. Amerin Society for Metals, Cleveland, OH. 2. WWdd, Hndbook. American Welding Society. NY. 3. SA E Drawing Stadard Manual. Society of Automotive Engineers, NY, 1970. 4. Heat Treatment. ASM Handbook. Vol. 2, American Society for Metals, Cleveland, OH.
172
"
~17-23
j
AMCP 706-202
APPLNDIX A
EXAMPLE OF A PRELIMINARY HEATING, COOLING, AND VENTILATION ANALYSIS
HEATING AND VENTILATION ANALYSIS The following is an example of a heating and vcntilating system. A-I
12. Kcat transfer aromw A so0.
Cockpit windshield Cockpit akin Cockpit floor Cabin rear ramp (uninsulated) Cabin windows Cabin floor Cabie walls (uninsulated) Cabin walls (insulated) Cabin ceiling (uninsulated) Cabin otiling (insulated)
A-.1I DESIGN REQUIREMENTS I. The heating system shall be capable of mvntaining a tempcrature of 60"F in occupied spaces where the outside air temperature is -65'F or above. 2. The ventilating system shall be capable of delivering no less than 2.25 lb of fresh air per min to each occupant. temperatureAAIof somn-. the ducts within 3. The .surface ... ......
50.0
50.0 60.0 26.7 225.0 180.0 259.0 105.0 225.0
6-.,in _y-Anc__t of
1800F.
A-i3 HEAT LOSS•S
A-i.2 DESIGN ASSUMPTONS I. The number of occupants in the aircraft is 33: 3
A-13.1 Cockpit The heat losses result from convection and infiltration, i.e.,
in thc cockpit and 30 in the cabin. Metabolic beat
rate for a sc.wted person (writing) is 400 Btu/hr per occupant. 2. Heat gain due to solar radiation at -65"F is negligible. 3. Outside air infiltration rates are: cockpit, 100 cfm. cfm; and cabin, 300 4. Humidity effect at -65"F is negligible. 5. The heating system heat loss is equivalent to 20 deg F. 6. The blower volume flow rate is a constant. 7. Mechanical heat sources and fan work are negli-
,a.
(A-1)
A-1.3.1.1 Conectim The heat loom are through the transparencies, uninsulated skin, and floor, i.e., " -ql, .,, + qj + qflw
,v.
a
wj
(A-2)
where -
gible.
8. The cabin ceiling and upper 3.5 ft of the side walls arc covered with 3 in. of insulation. The thermal coefficient U of the insulation is 0.G7 Btu/hr-*Fft'. 9. Heat transfer coefficients: Surface U. Btu/hr-*F-ft2 Transparent areas 1.69 Floor 0.7 Uninsulated wall 1.85 0.07 Insulated wall 10. Electri:,d equztpment uses 0.225 kVA "11. Air:o - t0lb/ftV, c, - 0.24 Btu/lb-eF
+ Qanm.
-
UA AT
(A-3)
(Btu/hr-*F-ft2 ) * (ft') * (*F) - Btu/hr
(l.69X50)[60 - (-65)] - 10.5W Bnu/kih q,t,,
-u,.,,dj (I.85X50)[60 - (-65)] - 11,560
Btu/hr qf, - (0.7)50)[60 - (-65)) - 4,380 Btu/hr
By Eq. A-2, the cockpit total convection heat loss is: Q,1,, ,0M 10,560 + 11,560 + 4.380 - 26.500 Btu/hr A-I
A-13.11 W9radm The •.. bhit lom soulding from infiltration is: SQ•,€ucWAT
( 6 tu/ib-tFj.
By Eq. A-6, the cabin total convection heat loss is •,..y•
(A-4) b/hr) - (-F) - Bwu/hr
-13,900 -
, :3,640 + 19,690 + 41.630
+ 2,270 + 24.280 + 1,970 109,380 Btu/hr
Basd on the given infiltration ratta of 100 cft into
the cocpit. the resultinS pounds of air W arc:
lb/hr
W - (ftl/min) * (lb/ft') * (min/hr) - (IOX0.1IX60) - 600 lb/hr
.. Theronwo by Eq. A-4:
Q~lw• " pWAT
18,000 Btu/hr
A-13.13 Toad Cdipk Hess Lw By Eq. A-1, the total heat Ions is -
A-3.2
A-4.
Since the infiltration rate is given ez 300 cfm into
Q40-1u,im - (0.24(600)[60 - (-65)1)
•
"The heat loss resulting from infiltration is by Eq.
the cabin, the rcsultint pounds of air W arc: W _ (ft 3/min) , (lb/fl1 ) 9(min/hr) lb/hr l -
(30XO.1X60)
-
1.800 lb/hr
'cretbre.by Eq. A-4
26,5W0 + 18.000 - 44,500 Btu/hr
-
(0.24X1,800)[60 - (-65)] 54,000 Btu/hr
C"
tion, i.e.,
A-I.3.L3 Total CaM.s Heat Low By Eq. A-S. the total heat loss is -
Celed A-1.3.I TIh heat lom ,.re tbroitgh the ramp, transparen-
cm floor, Wells, Ind coiling. i.e., Q
, S+ .: + qkwM,,V
109,380 + 54,000
-
163,380 Btu/hr
A-I.4 VENTILATING AIR REQUIRED
+ qaucy , (4.-6)
Baned ta Number of Clkuiat %W Mbamm Veuiladmg Rate The ventilatinS air requirements are:
A-I..I
+ qmaigwa eam OL" raw~
+ qkwbw',kw
Theaeoe,, by Eq. A-3 (I.85X60)[60 - (--65)] 13,900 Btu/hr - (1.69X26.7)[60 -- (-65)1 5,640 Btu/hr q.,- (0.7)(225)[60 - (-65)1 - 19,690 Btu/hr (1.85)X180)[60 - (-65)] - 41,630 Btu/hr - (0.07X259)[60 - (-65)] - 2,270 Btu/hr , (1.85)(105)[60 - (-65)] - 24,280 Btu/hr - (0.07X225)[60 - (-65)) • 1,970 Btu/hr -
A-2
WI• - W-,wkw + Wt.d,
(A-7)
The weight We of the air required is dctermined by: W. - (lb/min-occupant) e (occupant) - lb/min
(A-8)
Therefor, based on the given conditions of 3 cockpit and 30 cabin occupants each requiring 2.25 lb of fresh air per mrin Wa,, Wa.,,,
- (2.25X3) - 6.75 lb/min - (2.25)30) - 67.5 lb/nin
By Eq. A-7, the total minimum ventilation requirement is: Wa,,,,
-
6.75 + 67.5 - 74.25 Ib/min
V
--
_____ ____
706-202
____AMCP ____
vreqrurewut Baed an Maximam AlloUsbie Temtpeatume Diffeemg Since the furface temperature of %beduta wnnot cxcd 180'F, the maximum allowable temperature di•terence AT in occupied area; ;s: A-IA.2
A T - 180 - 60 - 120 dog F it ix necesaary to detrmin.-if thIs a;lcwable A T it sufficient to satisfy the cockpit and cabin heat losses
... 4•Thus -1
based ow, 'circulation demanded by minimum ventilation requiremenls.
or the a'lowable AT
Wa,,
-
120 d'g F is
MC.3 ib/min
- 25.8 + 94.S-
A-1.4.3 ToW Htalt Re,qlrgmss Since thc system heatin& loss is given as 20 dq F, the total temperature difference between the outside air (-65°F) and the heating ducts (180F) it: ATmw -
180
-
(-65) + 20
-
265 deg F
Accordingly, the total heat required is:
Qiow -= 0.24)[(120.3X60)J(265) - 459,100 Btu/hr
A-I.4.L1 Cockpit Requiremaet The required ccckpit AT is C,.kpa,,W/(cpWa.eg) degIF
(A-9)
Use the beat loss of 44.500 Btu/hr from par. A-1.3. ,1.3.
-44, 5l(O.2X6.A7X60)' - 458
,-'
F
'
-de4g' F tempatu-difface c-cg& '.nc th. allowable AT- 180 deg F, tim a*rflow to the cockpit musZ a.tceo the minimum required for ventilatiem. The required amount of cockpit air bawd on treallowabeA'Q 2 ,c -,t.•,/(c•.•7) alb/h .--1,545 l,5451bO/[(05.SXlbmi Ib/hr - 25g.8 Ib/min
• •€
... A24,, A-I.4.C•
~
R..eq(.'*
(A-10)
AI-.5
HEATER REQUIREMENTS
The net heat requirement to be supplied by the heater is the difference between thaOreqaired and that gained from the occupants and electrical equip ment, i.e.,
Q,.
A-I1S.5 Heat Ga6W The occupants and the electrical cquipmaait aft responsible for the heat gain, i.e., ,Q
-
-
m
"1
163,380/[(0.24X(67.5(60.j - 168 :deg F -
-,
(A-12)
The heat generated by the 33 occupnts based on tim given mnmabolic heat Tate of 400 Bt/hr is: (Btufhr-eccupant) * (occupant)
- Btu/fir
Similar'y, use the heat ks of 163,380 Btu/Ir from A-1.3.2.3.
j.,pir.
(A-t 1)
Oa, - 0.,u,
- (33X400) - 13.200 Btu/hr Since the electricd aystem power consumption is 0.225 kVA. the equivalent heat ir, (kVA) - (Btu/hr - kVA) - Btu/hr - (0.225X(3,413) - 761 Btu/hr -
Sinae the 168-dq F temperature diffrence also exceeds the allowable limit. the airflow to the cabin must exceed the maiur.um required for ventilation. TVi required amount of cabin air based on the allowable AT- 10 deg F is - 163,380/1(0.24X1:20)1
.. 4
'5670 Bu/h" - 94.5 Ib/miA-1.51 0 -- - 9Since
A-I.4.2.3 Total Air Require@e I \by Eq. A-7. the tqtal ventilation requirement based
By Eq. A-12, the total heat piaed is: Q~m, - 13.20D + 768 - 13,9EBftu/hr Not H Rq the total heat requirad by pat. A-l.4.3 was 459,100 Btu/hr. by Eq. A-lI the "stheat required is -
459,100 - 13,
-
445,132 Btu/hr
A-3
o.i.
A_
AW 7OWO2
,•
A-l33 Heater Size The nearest available heater size to satify the heating requirement of 445,132 Bt'j/hr is a 600,000-Btu/ hr heat".
•
A-I. BLOWER SIZE The blower mint provide air for both ventilation combustion :*: and .-. :For of fuel tc heat the air, i.e.. WaW - Wdk.
-v W.,.,,
(A-13)
A-IALI Veam. of Air t be Duhre The weight of voutilading air ftom par. A-1.4.2.3 is 121lbIb/min. Ta det, mine fh combuion airf'ow, assume: Heat transfe
S"1.
2. FI, Ikuflb
A-2L. DESIGN REQUIREMENTS The desip condiion is to maintain 90"F, 40% reuatve humidity (RH) maximum during a MILSTD-210 hot day (1036F and 95% RH) in the coctpit only.
A-2.2
DESIGN ASSUMPTIONS solar radiation effT•ts it is assumed that thec
helicopter is heading due south at 1400 hours in the
afternoon of August 1. Other considerations remain the st me as for the heating analysis. No heat effect due to mechanical sources will be considered. A-.3 DETERMINATION OF• EFFWFIVE TEMPERATURE DIFFERENCES
efficiency ; - 0.65
ASSOCIATED WITH VARIOUS SURFACES OF THE HELICOPTER
with higher heating value HHV - 18.400
3, Fud-air ratio W//W, - 1/15.
Tlwm put is:the airflow required for the 445,137 Btu/hr output iThe " (the •.- Qe/[(I1HHV(dW.W)i,] (A-14) ........ (Btu/hr)/[(Btu/lb)- (W W ]the (.lIb/hr - 54,132)/[(h 8.400X3 l/m 50.65) -558 Ib/hr - 9.3 Ib/mitz
A-2.3.1 .EioctiVn SeWa Temperstim outsde surface temperature isdependent upon outside heat transfer film coefficient fh the onortion a of volsi radiation I which is absorbed. and outzioe ambient temprrature To. The relationship for this efrective surface temperature is (solar temperture) as given in Ref. A-I is:
" ' • :to By Eq. A-1 3, *be total airflow required is: •:•••
.Ww•- 120.3 + 9.3 - 129.6 lb/min
~ ~or in terms of dininl• oriers
mf, Wa,,• - 129.6/p - 129.6/0.1 -
1,296dm
- T"o + nllfo. OF
ahe
(A-15)
fraction of solar radiation absorbed, dimen- .heattransfer coefficient (fdm), Btu/hr-ftl-*F
solar rudiation, Btu/hr-ft' I For this case, f - 13, To - 103. and a - 0.9; there-
fore, from Eq. A-15: A-1A Ptsure Dreg ?. Asume the followin pressure lousem: I. Per foot of straiht duc: 0.05-in. of watei S2.per 90-l elbow:. 1.0-in. of water S3. of water 4. Frsh AFr airtheinlz: eater:2.-in, 2.0-in. ofwater A bkowv with a prusst Aip rise of 13-in. of water is rmquired, applying the unit losmes.
A-2
COOLING AND VENTILATING
ANALYSIS
The following is an example of'a cooling and yen.ia a l . A-4
- 103 + O.81/13 - 103 + 0.061.0F
-,
(A-16)
The temperature for each surface of the hlichoptet depends upon thesun ingle, which affects 1.ThI following values of !h are taken from Ref. A-2 for the conditions of this example: Surface i, Btu/hr-ftl I vminnj. w1n,. V•;t INh. ise-,, V,,WW v.,,
IH,,,*,.• rTh I ,16
25
160
16
105 th
250
(
Um ?rom tbw I'.
kab oAw~ws Sq an Ok.mS frl *I
SqTsmpsvawL
Isis
Amma. UNMImmWe "w a"mum Of abs MOM*~q An$. to
A
IIM.S
Av
lie
A mw
40
Amm. tsmaip~rmo of tbe fo~owWS disms-
-T.Wdog F
AT-u
or I&
WW A'
(A-17)
4an- WE. a
-T..u
Cmakuom of tbe bW gui. bad onh do Sim dus UbI.
A... vm.
Tempamaur Diffuumm i 14.5 23
&T A Ts ATw
Amft 17.9 17.9 3.
A:, As,
The &s to be vied foe c.Iwlaionsam dolmrmine from Eq. A-17.
tATs
14 19.5
Arm,,
28
ATN
&Tw14
4.Ic~G. 7ue boat 5du. w Wa. uid ansmmum make UL., + fou
Q~qw
A-9
+~( fqAe.M
fo.n
AU2
lbs k (fkmfonr dboq
9
j~f+N
406 + 9'V + q(i + q,1nv
HEAT GAIIS Impnnnpal seen'su of beagh~oo lwc~midudas COMMPI
.
.A
-
2. Oncupewts
A-RA1
c....u.. ~ s.~
.~..
'~ma
)
~ ~ -i
~cA,i ibsthe dwctom.
(A-Il)
~ +jjW
qN
-(l.SIMM4) - 0
qs
-(IASX2M 93) -j (IJSMOM25 - 414
-721 DNm/br DBuyb
w S" +~831 +5S721 *414 2,W233W/hr
U
+~~ D,~,AlW/h
jgsi*s awm facingaukof tin AT,
(0.7)(Il 14) - WS DN/br A-5
A lam.
UwIlAP
yS.Abo) fter.w
orw ga
tvAr km~mt
(IX6I.)-
,917 k ./br
4. 3W/br
(Eq. A-
bfilhahme Cub 9A4L&.L2 Wy Eq. AA-4 adoth didp aummpim (.3W1p.IKEXUX13
-
-
owma
90)
A-U I.7
(176K35) - IAUflW/k
IAO 4-3.M
al It By ft. A-4S do tUd bea pis.
ADS
m bond as aTb.mle beot ame w o dro mn raws w 3sova#ae 2.25 t/ubs.uqnmm
1maip/b
The amous of watrrmifom kaw airsby is. *athRN ha M5to 4% amy be deswumims fy5r-o m a pugiwomatr donr. LA.. in 10301 95S 3j4 g - OMS lk.,w& r/~b-mi wrw in 9V. 40% RN mik - 0.12 b.-w /Urbmi
319 -3 O0w/k A41oa
To sm a amme" Nb Gai A44C d &Wa 11aftlas
aoftOs
waftr.
-
ThwvkW~~ ~ osldslrmdd ~ ~
T A-2A.L4
rinua a~tthesnu-
mad *a bast ouiavmuds
h Leg d O bbm his
Wemem(2253) -75 blob/m
(lISMS - 44 3mt/br
+ 2A86+
pmorarpb coadliloelg of the air for van&1*- 0&hi
lwaths, behi and uam, mmd prne,, qala whic opu~.i ast cmiulkwd in ordor to dateOw. fma mumu wAma Lbs ". of the air coadi&loel *ulpom~t us
Camilkioneg mad p o -ids
tI taauetmm A/1 ma S~. Tl. vamls fwam elr aditim 4 imPar. pwr. A-2A.1; the whias fwasl
6.9 am*3m/
is~I1765 ku/hr. A4WA4 ToWe Csshk Head Cisk s ~AI
The Qi.d.i.. frm pwr. A
ml. the dulp ,uqwirm" of 9F. 40~ RIL cow-
w A14
Qim-440
ldSa.
A4.S CSNub of Vaem" Air isat 1030F. 95% bha frshb I.ý to* ammapicbs iuwAaff LII. Skms watwim has WWl by oft
Gal Mw hu pis krm aciar miatm is by Eq. A-M. A44.L3 Odw Sm~m
-mwo (2X
l20u/br I=
A-13 ASK CONDMTOMU SZ5
quirwminb o pmL A-1.2 and A-22.2 gftuvwm
(3X4M0
Qmrauu-
md
Thhldar toaLd bean plid by ceowcil
aupeh
lvs asabaic bhat rmw of 40 Iku/br bwal ndw Skm
qs - (I.0)(179X143) - 439 ku/hg - (LI)(173X23 -~W6lii/br -
A4A.Z Omm~a Tb. WiS rinaw"te by the 3 wdkplt
wowe uIm Wi - OhG
Id
RMa
O *ahem of amidmom of uwa is1,07 Mu/ b-www ase tOw boa of ambamdes is
W%.
bwrAt/b-alr)
4e h
AMl 706-2 The heat removed in cooing the water vapor or water from 103° to 90F is very smdl and can bc ncglected. A-LS.2 Fm Si and Hoe1 Venillation fans and any other fans moving air into the occupied speac cool their own motors with the air they arm moving. Many timcs, ventilation fans amrinstalled in an exhaust mode. When this oocurs, no het p.ised by the occupied space and fan heat is not.c•ed. Air conJtloers may be d-.ven by compressor bleed air from eikm the cngine or an APU and ro "tmect bht is Wed to the occupied ospce. If the air conditioner it, driven electrically, beat will in gemeal have to be added t, the symtm. For theM am-
jt* Pis
pie cut. OMMUP
- A.
- It-/
- W -Jf/MD -(i/1*
ftl/hr
)
57,W Oft'/hr - (32)[(30X60)] - ,.•qi
..
p
(A-25) V ,(A P)/(33,000)y] - (ft 1/min) - (lb/ft1)/(ft-lb/hp-Lr)" hp - (63,921/60)(13X5.2)]/ ((33,000X0.S)] - 2.7 hp -
since l-in. of water - 5.2 lb/ft1 . The equivalent beat of the fan motor
thaeL
I. Vmtilmt~o fans are en the input side, bringibq fresh air into the cockpit at a rate of 6.75 lb/mmn. 2. V~,.dlatkm fans me moving the rat of thc 01cooled air at a speed ,- 30 ft/min. 3. Cros4uctiko area A for the cockpit is 32 fW ThI volume V.e •air to be moved i . ,V,
2. Per 90-df ulbow: 1-in. of water 3. Across fresh air inlet: 2-in. of wate 4. Across air oonditionwr. 2-in. of water. Assume further that thes unit loose result in a total lois in presure Ap - 13-in. of water; also. the fan motor efftiency V- 0.3. Th7 fan motor hornpower P is then
-
(2.7)(2,543) - 6,72 tu/hr
since I hp - 2,545 Btu/hr. A-2..&3 Tsm d Ralimwln Rqgpmi The total heat Ilcs to be rmoved from the cockpi to mtioý the design requirmea - ts pars. A-2A.4, A-2.5.I.
and A-25.J
QO*
In addition. the faflowiqamount f freh air intr•d,.edfo odu fo ed niaic(the dmiy p ofa at 1036F,
i
-
17,411 IWehr + 1.264 + 1360 + 6,872 - 39,W0 - 30,407112=0tio
3.3 too r-fripm-
""5% KH, is 0.64 lb/ft,) Summ 12,M00 ku/br
"V - w - ((lb/lmmn)
3 - K6.7U%0)( 1/0.064) - 6,325 f1 /hr
is
volum
ThU tal Vew
-
of ur to be iidlad by ths fan
57 •M+
"Auuemsm te fON I. Par foot of rat
,,,,
-
I to of tef.eratie.
(A-24) (min/br)]- [I/(lb/ft'))
6•n - 63,02f1 '/,& IPemeM dit: 0.00-i,. of wafe
IonIR
ZNCI
A-I. W. H. Soe mu@ad J. IR. FOdlws. Hefth aid AfrPsoml . Jobs Wiey asid ,1db Some, Iln., New York, NY. 19W. d. Ara& A-2. ASNAAE Hand k q, FMsW Refriptatioa. and AirncaSesisly of FlHeatin Inc.. New York. NY, oaudito•: fgiý 12.
•,\A-7
_
_
_AMCP
706-202
INDEX A Awasuibility in engine maintenwa . 3-I of gun installation, 14-3 of missle instalation. 14-7 Accesories APU. drives. 4-89 Cegine drive requirements, 3-15 trmanmision and drve systems, 4-8, 4-89 Accumulators, hydraulic. 9-19 Acoustic loadiug Set: Loading. acoustic Actuators flight control, hydraulic, additional requirements "-ind criterv 9-16, 9-17. 9-27, 9-28 hydraulic, 9-13 to 9-17 pseumatic, 9-43. 9-44 propeller, 5-6 valve, pneumatic, 9,40 Adhesive bonding, 7-12 to 7-15 --
puOCO
. Adhesives
inspection requirenets. 2-33
epoxy, 2-31
rdm, 2-30 nonstirctural, 2-32 phenolic. 2-3i proptstifs of, 2-26 structural, 2-30 to 2-32 Advisory lights, instrument, 10-3 Acrodasticity, total flight vehicle, M-83 Aging, of mulas, 17-17 Air conditioning, cockpit, example analysis. Appendix A Airfoil uctions
"antitorque rotor blade, 5-S1 "main rotor blade, 5-39 to 5-41 propeller blade. 5-68 Airframe structure analysis. 11-13 bulkheads, ; 1-6 cargo compak-w•nt, 1I-17 corrsion protecion, 11-6 com. 11-2 cns kIoarI. il-9 design and construction, 11-4 to 11-7 desig xonsidentions. I -I to 11-4 • " .eve et. I-.12 "e"ectrical bonding, 11-7 fatibut sensitivity. 11-2 11.5 11-S
maintainability considerations, 15-7 manufacturing. 11-12 material, 11-2 skin systems. 11-6 static loads, 11-7 stiffness and rigidity, I I-I substantiation. 11-12. iI-13 supports, i1-5 surface smoDthness. II survivability. 11-2. 11-4 testing, 11-13 transparcnt areas, 11-11 weight, I-I Air induction subsystem APU,'3-16, 3-17 engine, 3-6, 3-7 Airspeed indicator. 10-3 position error. 9-46 Alloys
aluminum, 2-4 heat treatment of, 2-5. 17-18 copper, 2-6 heat treatment of, 17-18 magnesium, 2.5 steel, 2-2 heat treatment of, 17-17 titanium. "2. heat treatment of. 17-I1 Alternators, design of. 7-6 Altimeters bsaometric. 10-3 encoding, 10-3 position error, critcriA. 9-45 Aluminum alloys See: Alloys, aluminum Aluminum Association. alloy designation system, 2-4 American Iron and Steel Institute (AISI) codes. 2-1 American Society for Testing and Matcrials (ASTM) nomenclature. 2-5
Ammunition ej.-iion of debris, 14-3, 14-11 feed, 14-6 onboard storage, 14-5 Analysis example: heating, cooling, and ventilation. Appendix A
gtear designs. 4-35 to 4-46 geatbox housings and cases. 4-64 rotor brake. 4.63 Spline, strength of, 4-39 I-I
~~~~INDEX (Couti-ed)dun 37,-S
AMCP Analysis (Cont'd) vulnearbility. 14-13 Annuuling. description, 17-17 Anodizng, 2-35 Antennas avionic (prwa1), 8-10 commu'dcation, 8-13 high -fraquencv. 8-13 installation of, 8-12 low-frequerk-y. 8-13 UHF, 944 VHF, 8-14 Anthropanwetn' 1&is. in. cockpit domil. Anticillision )*Mhs, 13-9
43-1
of a*,"~ ti. indwb'f-;On Ssutrul. S'7 of pitot-static system, 94i7. 948 Antdorque rotor See: RiW system, antitiuoque Arniarnen boresihfins, I"contrcis (amnirkj, fuuia, uauie rdk.-aw 10-4 synb~im. dessin Suidenr%, 14-1 to 14-7 ~~iy~ ArMOr ki~tcrw torso. 14-1btO@
atutacbnit of, 144~9 itslm~ 1'5 &.veoqswt or camwufor. to imnubitiation oý 10ovift Paus by d"sa 14-1.9 it.6igenoii, 14-18 inastllati,-a, design couidermtioa, 14-18 to 14-2D inut4gl 14418 materials. 2-27 1ýp 2-30 s.;-Jionof14-16 parastic.
4-;st*ivii,
protectioni or tranamission aztd drive syssm =xtponents, 4-23 removability of, 14-18 Ar-iiculsted rotors blade rdwt-~ow for, 5-27 to 5-29 &4-e .bo: Rotor sysem a'JIMu.Roo swa, maintypesand knessau Assembley, rotor blades, saatve, 5-%6 to 5-30 Automatic dirsction finder (A[W) 8-S Automnsti latoequwpmmit, foL maiaacvaommm 15-2. 15-3 Autorotation entry. stability .equirhakawt. 6-8 simukation of, 643 Auxiliary powvzr anit (A?1U) aem drives, 449 blesl t'it type, 3-1S contf-1s. J-11 cootig requir'nnent. 3-1S -Sto 3-7.2 des4n and innallwi on rqimmn. 1-2
ecatne Muwtia& by. 9-3 wtn.31,31 ehu falur. made mod ~fe anals-sk (FIEA). 3.21 sysitun 3-19 Wuo irld air dactiag 3-46 lubcalim suson. 3-20 1isinii" 3-21 wi~afteieace faounilsq sysbun. 3-IS pneumsatic po&muskiws, warupncy 9-33 protecive devices 3-18 Mitbility. 3-21 sc-lty provisolo, 3-21 "tatW 3-20, 9-3 Mimic subsytm cUWA;Aunicatim.~ I!-S. 8-4 deulg acohws~tiaoss. -~ 6-2,89-10 Offict G! MWo MO&Wkmti ekxtrw~agnstic conWrtlbulky. 64 8-2 tam 5 -2. S-? envnonmaitai A.,a ootr~, 1.7 to 6- 13 M~iuresseas, 4-1 $on" ines. M4l 15.6 U04-4 to 6-? "P 91pM and track, -,owo blade, anta"Mq~e rotor A. dri~v shbufti 4-7 propd11ar, 5-73 Ball betri4s~ o.t3 sw.ei 16-10
to54
L)P5S, 114-3 in 5-31 bt, in rot:.r-badt SL-iitic twt~avwe, mgmerictkea' *%No&ug DUa wotA. proc~prtie of, 2-24 110taies chart- ri 7-576 chuing,7-.7147 i'stalation, 7-27 lead-ncid, 7-1§. 7-16 tnaitenance. 7-1S nkkt1 cadmium,. 7-15. 74b~ silver-tiuc 7-15. 7-16
4-76
utilimatica~ kLds. 7-16. 7-18
Betring lift, wlculatiots af. 4-53. 4-45, 164 to Mt-8
beari-q aska typ". 16-15 beaings aitfraani, WA-2 angulsr cont&&& 16-9 aantifticdw. cleagiflktiou of, I1q-6. W7bultys of, I"-
bag
(Cesd7-d
(WDEXd
OWLut pulctlewlo,14.47
dwk. iecic. 2d m 1.4to16-16
drive
elmiomawn. bwaw
wA
440oaks
16-14
in raw~ blede rounjem,
inetallation. 16-2
S.3= 5-33
wea, 16-2, f7-11
pssbwm. 4-42 to4.57 draw1@ .4-3% falgi. Nk ".45, 4-56 internul cherSUNtdIVM 4-S1 to 4-54
Banded WLucture am~oOy, 17415 4wip,.2-33 Bowling adhusiw
maegimma4A to 4.50
d
Opuraliq wikhom hbdcudo, 44 skd~ iNbval, 4-A4 4.53
11104whAu
ftls I461
boetstiiuk-lbive 6-14. 9-2
I" 3- to 134
bDom fimoeano,
smoneu rod..td. hbI' e 93 S mINu 16-10.1I6-11 pitb W~k. 6-21 po" lows it. to &%aim~ 44 uuaistiono. W1 4IS raogr, typs ot 16-10 to 16-12 roaft Owew, hn rowo Made. weieaiems 5-3,
9 T~cw fabrik in rwot paws. hm&4. ULk
pr
blud ionhomi
5-31
~M~tt~ŽU
oo~wd wo"6u 6-? traeniomni sad drve qviewa, 4-29 to-3,2 Mock oma4. for coating to -im -- M , 2-36 bal~m, sa" track. rw. 5-46 to 5-30 WAUcadsk.Moijg u 5-45 p c~ *damogi ~~i
..-
r,5-~w. S -70 o5.1
5-41 to 546 No& ifobdia& S-3J so 5-37 nawo,
SA&l OMdsP redlv. ofasry S-4 I"ud MMtele pvupdhv. 5-74. 5-75 w-. MO(to 3.53 WAdIS OphM~s rotor 5.36 Medursemoions
pr~dl. 545 5-46~ucwr1
L\
2-1S
rowor by*audic, 4V6 to 4-41 9M whuul, 9-&.9-2k,124 baziv&g 17-8. 17-9 bsi -e~q~u (BITE). 15-1 bolhowids, 11-6 Bullt *0e and *kdk, 14-30 11" liailers. pa. 14-11
C
20cow"e
$41% ~uaby. 134
I
scm uhie,2-31. 1742 to 1-15
lcriablam 11-7 boesmeycb structarva, 17-13. 17.14
sldonwi MahWqmS Meete, 5-SI PVGPeller 3-41 tuKr 3-36 Blade br atau balnc med amh. 5-46 to 5-30
Wf~hSM ies"o systo-, &23 jeaeag.17-I5, 17-16 struivwl. 16-31to136-33 C~dadmi a p1m 2-?,.2-36 curp ( wrkw extera-il 13-14 to 33-M "urmL, 13-11 a ~ y o u ~ m~ m 3 duig c-molderatiews 11-17.123-11 foRm. 114L1$I-1I,13-13 Carg dier mad ramp du.k I 3-IlI hydraulc umwmyv~mm 9.3 Canto. lextrmd, vouncmm syqnsew 13-1w to 1;4 Cargo Waeding, provimieme far, 13-13
CArgoiedowa load-l~Inadg, 11-9 prvuiem
413I. 13-12
alumaiimun al"i~n.m2-4 ym o.f 17-1. 17-2 Ca"06ow Oimh. 10-2 cbmewim piamiog 2-36 COrak brahms hydraulic, 4
I
-
CAXUaM mpa to u=twml dbusajs,^ S7 to pftbt.6.7
Caps, ums of, 16.4 Ca".g @iiukiuugg440 t.442 Wrk& 4Q
o1ro
11,3 -1.mI to.1 6.
awimic, 2-M4 psp.,2-36 aminW*aW,
813-
fkug . 6.15
w I
apomwý 84 AC I.DC 714 Dc to AC, 7Ms mul...
~~k~y i.hi,154 UHiF/VHF e..pws 6-3
camp" am ma.
&k, ws.
AC umsu.m 7-7 AR) mownd 3.18 ookpk md c@W 13.7 13uh Appu A amaqi
4-76ato4" h A 827 imuh 9X9.3 bywha~ pwoo" gmoms 9-33 to $44 jpulwu.i 545 to 546 *frE
cmod
uww~m
as*)
umlss,
"Swa
cow~w~ut
taia 241 I2.t
- ~~
pkhag 2.36 conW.WNk uA*, &Aammiudsaw. 3-23
~ 246
&AM li p-mh i. G xoou,,
.
441 wo 4417
dhs%34
Ukmomi 2-M9 2-30
of
44 w 447 Wm.
~
w
~~
9mk 9-3%
Ccgapu drift aik.
818 cmw,
hymik,9I9.^ Be25 rini rwdh hmw.NeW cwaom. b"% SowNAal oudm Cowmur *AnWwt. gi am *Una. 34-l pkamai. UK~ fcDIL D" a cgafsut Cow" =wag*~i. i~www", G617 &t:.not ww bf~mo S-al "-*. b4v kaow wltiWhh. v"mi"a 6-17 kmainýwr. 6(44 oa m.q*644r mohm Co" aepm~u oqMPAW &IT. 1.4
-I~.55po ,iutý-32. 16-33 ~
1-1
4-76 w 44
.4
lpmn~w oad aCP6 umo, l3.. clubwveiwo u.k 11.4 0 -imm"in ia. f"*am dmWa, 3.10 Vww6 uI-dFAw eo.M 'm~bWAICM 154 MbCCukpl Sft~' CN~ *" 4 IN. 4-74 9"*hd
-,
WIpawJ. s4a 10 S412 t1451
r
_
_
_
_
_
_
_
_
_
_
'Mc
_
7W~2m
INDEX ((Lontdme") Elauiomewww couplin& drive miaft. 4-78 Elastomneric materils, 2-10It Electrical oinknctk * ito hydraulic coatporcuss,
Veers. 2-9 Damps'u WC "xkm. 5A4
Ow *AU49 -WOS mWOmui SWPWn4 641
9-29
weduiotim reskmmm%, 6.3 to 64t Dein dimposo
Wammumic a,
14-1l
IlAggMO/hu**M wW*"Pshle, ID~imseua syusm requireamutt fot, 3-11 Duip pwanawus, rotor Wiade, 5-37 to -5-41 Duulg reviews 1-1. 6-28 DesectAbilt drive sysms so.., contributiom to, 4-17 Dia4wekt Istsniqucs frawnijsso armd drive %y3to.. 4-70 DirmtioWd comerols 13-3 Dimzlaala rasetale bob&Ig of, 2.32 brazing of. 17-9 definition of. 2-7 use o(bolm in. 17-11 Distane trmu~ringi equ4mai~en (DM13), 9-3 Dopplas radar s)sAwxm, M/
tI9,
DmwiE gearbox bearins,. 4.56 gearbox hoawiip iwd cues. 4-67 Pars. 4-47, 4.48 sow&ne 4-W) Drive piad., gtboyr, ammiraim, 440D Drive slJhiý 4-72 to 4-81 Driv.4 wyime. general a kruir.matu, 4.3 Vipe
S.-1
I'~5iu~n.
in Oýtrkia! rytmn. des*gn 7-21 1.A Cio"okyftra1 Serve VidVe fromp. 9-27
1-2.2
Ce3-13
Ormos
Electrical eupment~. invAialation. 7-24 Electrk-cal power exiternal. 15-4 selection of sou"?A 7-1 Electrica suiusystem batterims 7.15 to 7-18 clactromnagsietic interference, 7-21 to 7-24 fittings, 7-29 geneal requiremoiw, 7-1 to 7-4 generators and motars, 7-4 to 7-15 installation, T-24 to~ 7-27 lightning andi static cl~trcity. 7-29 to 10-2 maintsinability conskicrations, 13-6 ovccload protecion, 7-19 to0 7-21 voltage regulation and revee current rciay, 7-1S. 7-19 wirc 7.27 to 7-29 Elcmamagi~k.conpatibility (EMC), in aviceiz syrs-
Of~t.nw
Du~t tolL. 6-,6# uyowa.~c toods
exterewa cargo, 13418 grun insaallatioos. 14-4 Dywauiics witor~vc tolmi. 542 J~J(~~ r47to56
w.I~r 5-16 to 5-27 Etlgas, 2-19
effec 0ý 00M. pitc% up, 41; geArbox, askawmsi of, 4-30, 4.31 t) rnmo sn siemttem, 4-4 EýýI. amus, C'*wOCAimgby. 3.4, 3-14 -ka~eram XrBaigea~aei
zupn*-a*,rjof, 7-22 Electroa bftm weling (EDW). 17-7 Electronjic countermieasures (ECM) devices. 8-7 Emargency drvice, pcesaunzc componeots. 9-33 EnhCrgeac liHS1ng. Passene COMpartenwrt. 13-7 Etncy labricat on guarbx "47 taounimilon 41.14 .a&imr wy'11 4.22 f-a railv% d-toul cuhirs6,* 1&.29 Enihulc limit icar tecth, 4-46ý -
-
rtoritamwtic m~mah-n adudius &dtnc"cctPosWna, 5-55 laing ptop~mr bMadcm.. S-76 rotor bwiov., 5-S4 to 5-57 Engin -'ss~y 'liW, m~uiftmwat.n 3-15 Engitz alt inshmclic~ ssahe*atam 3-6, .107 Eai
wWN tJyuain,
3-13
dr1p clwekj.'-1. 3-4 dsit,. 3-5
typles o.3-.1
to 3-ti
F~.I urcto ap~~ka et
4 -
INDEX (Condami) *Engh
ibine
n ito f.35Ftgefv nyustm. I-SEqangieeig pbasims 24 Eavuomenetal coanvW systsms. cockpit mnd cabin. example aoals*s Appendi A Epoxies as stutural adhesives. 2-31 deacrlption, 2.12 Equations of motion. in Viability analysis, 6.4 Erowio protecion, rotor blWad. 5-44 to -542 Exhaust Oectors, 34k 3-14 Exhaust subsystem AMU. 3-17. 3-18 engine 3-7 to 3-9 Exhaust supprmeomo 3-9 Explosivt fornuins metals 17-3 mocurriaeof, 3-14 to 13-20 Extunions m".al 17-2
*=Ensigfsswh
kVExuraW
*F
Fabrktionshop roceses. 1-3 Fabratin. rocon, sop 7-3bydraulhmsyv=6 Failwe mode and eAMc analyis (FMEA) Y5 P'~'Pd.
*
S-~appl~catioas.
stabil~ity auum~awtion "yowem 6.12 weapon syseams 14-10 Failure mod%* overrusning dutch. 4461 rotor brake, 442
Primay.4.32to4-34 secoodsry4--u rats.a trauaission sysmem 4-13 to 4-1s F#ailres par tooth ~4tt~. 4-0 ~awie. 46 swxiag. 4-41 to 444Fre. Failures, red~wason of scunas. to inrasing r&&"blty (hydrauklwvtsystem) 9-5 Fan
A *Fauihre
Appeadix A \xA
ammesby
5ooc 43 composute mauurials. 5-55 ero erns 4-55 propelle blaW, S-75 to 3-77 rotor system coupomemla S-IS, 5-53 to 3-57 Fatgue sensiivity, in airerme dwkin, 11-2 Fatgue tesin Mtudwar members, rot oruee, 5-56 Feibadr, for actwauor pods.n csanuL, 9.17 Farou ustal applicatos aow attrib~se 2-1 t 2-4 bea trhamarns 17-17 Fibugiss laminates, 2-11 Ci~le caps%AeL, 3-1l k,-,ters APU. 3-17 *Metoic, for EMI protaitos, 7-24 3-13 snow ffiI syst, mbox krabricaon systm, 443 s9-21. 9-23, 9-29 UMes saU , -3 Filtration, byJrauk sygtm. 9-10. 9-11 antitorqus
2-36C Fam control system. 8-7 to 3.9 Faae detorsi At,3 y o,3Fin dwewmis MWNS APU, 3-21 -sme copctwt 34.13-7 Firuwal MV -19 3-5 Iui o3Fus i OqWpowset lw wwse kw 13.7 WR arrmISoecrical, 7-29, 16-16 to 16-22 b ~ 2 ,I.4 types of ro
quack-viWeak. 163eauinas threaded apfmictkaw 16-, typos of. 17-10 to 17-12
-9 to 51
5-35 16-2
Fah y actuator (by**Akmi), 9.13 to 9.17 dsg md6.~ lad__it62
tosht~ 5-2356 14
.
2-34 to 2-38
moddertlsus
to 6-26
I
) i
AM~P 70M20
.
~INDEX
(Coadmued)
Plight cusMW systM (Cone'd) latesystam uwithin (hydmau~lc) 9-8 kiemanac ulhcsa, 6-14 p~antlaaily ca uldustiona, 6-25, 1355 andonrotaing) 6-is to iwxbnismms (robtlvS= 6-26 nonrotatlag, requiremnats and design gtandards, 6.22 to6-25 Pilo effort rMqureMefts and consideratioms 6.14 to 6-18 redundant hydraulic power sources required, &-I, 6-18.9-1,965 relisilty couulderutie. 6-1S stabilty and constrol raquirunee and specificstlooks 6-2 to 6-9 usabiiy aamVmnUAiWo sYOtM (SAS), 6-9 to &-14 strutura salthus. 6-24. 9-16 teing. 6-26, 6-27 vibration feedback 6.14 vusablunifty amsdralos 6-18 Fligh ingxnuaints, 10.3 is
~
~
4-74
idrie hftdn
Galvanic action umme~ebf~ty of mink to. 2-7
plopulr 5-76 Flig
"ht eeng in
in-
-~~kOof fliht contol
1.6.npneumatic,
maw.
P-P 1sw i Me
5~limitilees Swe SaM flmi
cojPly-by-wiwe. an FWMla qasahs develop=Me of. 6-2
hydranhc. 9-30 %-41
Gearbooss, lubrication amd woMMng 4"I Glear =Wsqzngu% 4-78 Gw loui frictIon 45, 4-71to 4.11 wind* 4,.34-
FWlVefdtha 11-10 Fltstiim mmipacy *mpbflty, 12-10 6@4k ~124 Plotte andivAmn 543
-n~r
Fuel controls APU. 3-19 cockpit. 3-13 Fuel drain valve requireina s. 3-11, 3-13 Fuel dumping provisions, requiremnats, 3-13 Fuel gaging systeaw 3-1l Fuel subeystemn components of, 3-10 to 3-13 crashworthiness caiterla, 3-10 definition, 3-9, 3-10 diagram required. 3-13 drains. 3-13 testing, 3.13 Fuel tanks expansion space, 3-11 external, 3-11 integral, 3-10 refusling and defudlIng 3-11 vents., -Il Fuses electrical. 7.20
Oa sstm
4-35 typws 4A4.4
6-34 "rol, analysi of bondigfat sigu urungtk, 4-36 to "I4 plief" fellers 4-44 to 44 "I 4s4 waION failse .4
uigai&mug raaMhUa 6.2 Foaf r* gjpropor" of, 2-23
OwMa
Folding Amawaft roci (wAft), 144 Fodn.rater WeS. Sue Uska WW"in Fworgn motnk, 17-2 Pornade lighW. 13-9 Formb&g am"n Mork ds at~17719 .. *mushodeof*dwn4 177-3
designw ansimalsia 4-34 t 446 draw apned qui&Meat &,4.47.44 01sawaan AC, charectuidMic of. 7-6 Dc. chuac" * -" of. 7,. 7-10 ,7 74rei vuspowm fat~ua~ mpsar"hposen.--
Praww aifz s, = ns tonkual I I-s Freqemx7 mien, rowo isisewk 35-19 to 5.24
)
lopwarais.o
in akdus. dedp, IlI-I
Friction OPMstiftse to pwbox N'P' P..4&4,47
Geunuatry roto wlaes 5mua -37 to.541 Obbal roWo. bb6 meelI to -*5-29, 5.30 Gmrlf h re PbgiC (OM.P fabriMeje. mihade, 2-13
2-12 to2-16 4VMsmtys GIns Ol~ mi'iohb -1O 1-7
AMWP 71-202 INDEX (Cotdnewed) Honeycomb struct',.s Governors, propeller. 5-66 composite materials, 2-2D, 2-21 Graphite, 2-18 use of adhesive bonding, 17-13. 17-14 Greases, application for, 2-38 Hook. external cargo, 13-19 Grommets, uses, 16-4 Hoses, tos in hydraulic system, 9-12, 9-25, 9-30 Ground resonance Human factors avoidance of, 12-11 in armament controls, 10-8 landing tear design criteria, 12-1 in cockpit design. 13-2 rotor design criteria, 5-19 to 5-23 in maintenance, 15-3 Ground support equipment (GSE) Hydraulic circuit breakers design considerations, 15-1 use to improve survivability, 9-8 standardization of interface. 15-2 Hydraulic components Grounding, electrical, provisions for, 15-4 electrical connection to, 9-29 Guided missiles See also: Hydraulic subsystem, components design considerations for Hydraulic fittings See: Hydraulic subsystem. installations, 14-6 to 14-12 fittings See a/so: Missiles Hydraulic fluid Gun indication of level, 9-26 accessibility. 14-3 location af level indicator, 9-25 burst limiter 14-11 selection of, 9-6, 9-10 feed mechanisms. 14-2 uses of, 2-39, 2-40 fire interrupters required, 14-10 Hydraulic pumpM, 9-18. 9-19. 9-30 locations. design contiderations, 14-2 to 14-4 Hydraulic reservoir pintle-mounted. 14-5 adda6*,oal criteria, 9-31 pid=moun-*. 14-4 design requirements, 9-20 turret-mounted, 14-4 subsystem Hydraulic types of, 14-1. 14-2 accommodation of reLative motion. 9-12 Gust loads analysis required (beat load, mission profile. effect on propelkr vibration, 5-62 peak power, and total eneay). 9-12 rotor systaem, 5-24 to 5-26 APU and/or engine starting, 9-12 cargo and/or personnol hoist, 9.4 H cargo ramp and door operation, 9-3 Handling qualities components evaluation of, 6-2 access for removal, 9-25 methods of analysis. 6-3 _agmput•m
I64-!6 6-;.
u,_",,n and &. .,An011
0 _1
Se aW Flyin" qualites Hangers, drive shat 4-75 Harnesses, safety, dwe of, 13-4 How eicbnm
control scector valves. 9-22 daidgn and development data and reports, 9-13 design *oside tio 9-7 to 9-13 desrgn pressu, 9-6
S amso: Cooler metasb Het rummmt descriptiou. 17-16 to 17-19 design coamideratios, 17-19 Heaers, groud 154 it anld cabi 13-7 Hetin. exsaipls amalyis, Appendi& A Higlsaw rotor. blade mustioa, 5-30 Hinge autijeimld rosw, 5-9 aieslud (vseerW mrtor, 5-10
fittinis. 9-11. 9-30 flight conol pvwtr. 9-I. 9-2 fluid W. intcoato, 9-20 bat exAnu, 9-13 instalation comsraximm 9-25 to 0-27 maintaimbdity a dwatkmhk 1,-.7 sisollaeuus disipju acritu, 9-77 ,a 9-32 operating p'auune casidpaio, '.P-0 Pressumr r~ ltioa, 9-20 propelle c"Wro. 54o~ p-'stioa attumatikos 9-19
Hoig
r.iabilit. 9-S. 9i4
bydraulic e---at- mi for, 94 intearal, cao oding. 13-14 1.4
roam brake. 9-4 atret criteria, 9-6
dzg
4-86
K
Ilradtmi,•09-10.91
I
INDEX (Continued) Hydraulic subsystem (Contd) temptture cni
tion M9
vibration, 10-10 9-.7
utility sys'.un. 9-2 to 9-5 wheA brakes, 9-3 Hydrofoimui& metals, 17-3 Hydropn embdttdment, methods for ,diuf of. 2-36, 2-37 SI Indicators adviaMry, tramimion and drive qsytem, 4-70 attitude, 10-3 attitude dIuctor, 10-3. 10-4 coure,
1()65
dlftru Prum, bydrmalic filter, 9-21, 9-25 dfhotivemm., pomumatic system dehydrator, 9-35 fluid lkeve, hydraulkc reurvoir, 9-25 horizontal situatios 10-5. 10-6 P.,,Sue emapc air bottk 9-33
bydahlic system, 9-22, 9-30 p wnmat sygam. 941 rulio magnetic (RMI), 10-5 go"fof 0MV, IUj-3 •.atic paund-ank 1'09 .trn and bank. 10-3 vartical si tion. 10-3 wutq, parking brake, 9-S ;•.
€
,,nzi,
systems, 10-.4 m Systes, 101 10-9 10-98-4 1 weepon • na,&*atin Inerial o Infrared (IR) s APU exhausit 3-18 exhaaust. 34, 3-9 quireamunt (vsle)k. 3-3 Inlet air &xatiAg, APIU, 3-16, 3-17 int wSdliw prot*cn of, 3-6 •IinIPEI•o d in fia e ld 15-3 trammiaWn and drive system, 4-10 i~alStahI ea ocal, 53-23 antitorqVS rtor, 543 llap.a& 1-14 s•emnLal See. Grou( d rmonaonce -
•',
. 5-14 p.hdA-l 5-13
tlanumatla a-26. S-27 :iNrMunt WMAWi sysm (ILu), ,• 5 I u a sAyOte S
amibdity
pealKI
"inteh
of oospDMPMa
MpQMMt&
AMCP 706-202,
, 10-19
?.0-1
consi-ations, 10-10 SMik% Mquirsmtat. 10-1 Msmainsbhity consiidurations 15-7
warning, caution, and advisory signals. 9-2. 9-3
Instruments flight, 10-3 to 10-7 helicopter subsystem, 10-7 light emitting, 10-1 navilation. 10-3 to 10-7 types, 10-9 weapon system, 10-7 to 10-9 Insulation of electrical wire. 7-27 sound proofing, transmission anid drive system, 4-1! Intake screen, APU, 3-17 Interchangeability requirements. blade balance and track, 5-46 to 5-50 Inter•oammunication selector, system, 8-4 Interfaces, ground support equipment, 15-1 Investment L.uting. 17-2 Isolation. vibration avionic equipment, 8-3 engine. 3-5 gearbox and drive system conponmAts, 4-12 instrument panel, 10-10 JUttisoning gun and ammunition, 14 3 gun pcd, 14-4 missile launcher, 14-7
Joining. metals, methods of, 17-6 to 17-11 K Kinematics, rotor system, 5-7 to 5-16 L Lacquers, 2-35 Lsminates fabric, 2-16 high pressure, 2-17 industrial, 2-16 Landing gear avoidance of ground and resonance, 12-11 bear paw. 12-9 components, 12-3 load analysis, 12-11 maintainability considerations, 15-8 retractable, 12-9 ski, 12-9
skid-type, 12-8 water landing. 12-1, 12-12 wheel, 12-1 Landing loads, analysis of, 12-11 Landing/taxi lights. 13-10
Launchers missile, 14-6 to 14-8 rocket, 14-8 to 14-12
AMCP 706- 202__
__
_
_
_
_
INDEX (Couthmed) Life bcarings, 4-55, 4-36 tronsmission and drive system, testing, 4-32 Life rafts, rcquiremcnts for. 13-7 ,Lgting exterior, B39 instvument, intensity watrol, iO-1, 10-2 ins'umentation, 10-1 interior, 13-10 p, oteation dipxtricdl nyste,, 7-29, ')-30 rotor blades, 5.D3 Lightadviiory, caution, and warn~ing, 10-2, 10-3 anticol1ion, 13-9 cabin and passmiger compartment, 13-10 cargo compaim.nimil 13-11 cockpit, 13-10 entargency, ?3-10 ,/fotV(atiC., 13-9 '!landi.!tt~ki- 13-10 panel, i-1,'i3-;O portablr, insjection, 13-10 posltion, 13-10 troop jump signsi, 13-11 "Lirait-cycle oscillations, 6-7 Litt",s, instalhitio "oviaions, 13-5, 13-6 [Load analysis battery utilizatio%, 7.l' electrical, 1-2 to 7-4 larding, 12- 1 LoaJd f..tors
1gearbox
iLightning
I!,
g.ust, 3-24 to 5-27 tires, 12-4 Load-limitis, cargo tiedo'wn. 11-9 Loarling aco-sti3 rotor, 5-24 -anli~oque rotor, 5-82, 5-83 "pr.)pellerbut, and blade rlention. 5-65. 5-66 L'3ad~ig ramp, c(argo, 13-13 I •Loads
gcarboA housin~gs and -cases, design, 4"64 to 4-66 gust Se,: Gust loads whzel, Cu ca.go compartment floor, 11-8 "Lougerons,) 1-6 Loirg-range navij.tion (LORAN) equipmenm, 8-6 W,. transmission power, 4-5
4
"Low-fuer warning 3-11, 10-1 ,'"irnts .•.-fitnt, 2-39 filnt, thicknesm af, 4-7
"t
r onwt.tnt., 2-34 I-jO
Lubrication boundary layer. 48 data lit required. 2-38 drive shft beoring, 440 emergency, turatasima mad drive system. 447 gearbox bearings, 4-50. 4-51 hydraulic uyatem cop•ponmeni, 9-32 nregmes of, 4-7 to 4-11 roging element bearinp in blade relttlons. 5-30, 5-31 Lubrication subsystem APU, ,•-.0 engine, 3-14 transmissOn and drivs, 4-41 to 4-88 M Machining toicrances, 17-5 types of, 17-4 to 17-6 Magnesium alloys casting, 2.6 dcK•=riPvoio, 2-5 limitations. 2-6 mracbinability, 2-6 Maintenance. desigri . siderations 15-2 accessibiiiiL, I'-men engineering, 15-3 insmcction, test, and diapniis, 15-Z, safaty, 15-1 •tandardization, 15-2 Mainttiance and GSE interfaces airframe structure, 15-7 gor, oroan~aIt, andf prteb A-ive itystemn 15 avionic subsystems, 15-6 crew stations, .15-8 electriCol subsystems, 15-6 flight controls, 15-5 hydraulic and pneumatic subsystems, 15-6 instrumcAtation subsystems, 15-7 landirg gear sLbsystcm, 15-8 propulsion subsystem, 15-3 rotors and propellerm, 15-5
transmissions and drives, 15-S Maneuver stability. criterion, 6-7 Maneuvers, effect on propeller vibratory loads, 5.62 Ma,-ufadturing, considerations in airfram: design, 11-12 Map :-ae4s, location, 13-5 Map d&%play projectrd, 10-7 roller, 10-7 Marling eand painting, helicopw 2-37
k s
AAicM"W ý a INDEX (Cointinme)
Master Wad"s
N
for propeller bul-ncet 5-73 for rotor balanc and trac, 5-46 to 5-50 Matchmed die mklding, G3?, 2-14 Materials airframne structure. 11-2, 11-3 armor, 2-27 to 2-30 composite 2-li to 2-27 drive shafting. 441 elastomeric. 2-10 gear, effect on fatigue strength, 4-38 gearbox came and housings, 4-66. 4-67 glazing. 2A0f hydraulic actuators 94~6 main rotor blad., 5<-50 to 5-53 metallic, 2-1 to 2-6 nonmetaLlic, 2-7 propell blades, 5.74
rotor brakes, 4.64 rotor system, fatigue life considerations in selectiou of. 5-54
tharmop!=azc 241
42Wt
thermosattin& 2-9 Mechanical instability Swc Ground resonance Mechanical properties sources 2-1 metals dassaimlar, 2-7 extrusion of, 172
feru.21Io2
fabrication of. 17-3 frigof, 17-2 nt f~q~iw 17.1trarit-ninsion joining of. 17-6 to 17-16
'plating
.4,Missile
S'
Natural freiquencoes antitorquew rotor, 542 propeller, S-60 rotor, 5-19 Navigational equipumet installation considerationn. 8-4 maintainability considera-ions, I 5-t Noedle roller bearings, use of, 16- kI Nickel plating. 2-36 Noise analysis for diagtiosis, 4-74 gearbox, effect of gear type and tooih p'~ch, 4-12 gun installations. 14-l1 transmiosion and drive system, 4-11 Nonferrous metals description. 2-4 to 2-7 heat treatment of alloys, 17-18
Nonmetallic materials, 2-7I Nonrotating flight controls, design. 6-22 to 6-25 Normalizing, in heat trea ment, 17-17
Notch factors, applied during esidurance limit testino, !15S Nu~ts fixed, 16-2 nonfixed, 16-2 ussof. 16-2. 17-11
A
0
spestrographic analysis or. 4-71 svstem luboricating, 4-85 Oil tank and cooler, APU, 3-2.7
nonferrous, 2-4 to 2-7 paraatmperiaw to9*x n o, 21, -4 of, 2-36, 2-37 work hardenir4 of. 17-19. 17-20 Metalworkhv~. description or proc~aes, 17- 1 to 17-4 Micropoon.-beadset. desig2 requirewents. 8-4P launchers, 14-6. 14-7 Missiles, gsuidid, installation. 1446 to 144.2 MCck.Up. revicir Wand "Valuatioii. 1-I Molyldiwo-- disult.a as lubrikcant, 2.39 Motors eacticaI. 7-13 Mounting Lystzi.i. tcnamia and drives, 4.72
wiltr 1abSee: Vibratory loads utp.-. 'ac4 ntrlsAPIJ, 3-18 Ovcr'.,,d prut.rction, electrical systrmns, 7-19 to 7-2: Overpg, wer utesing. gearboxes, 4-~1w
huzzle 10ast, miinsirtaition of effocts of. 14-3 M"~
Perfocmawice awd.al'ss tLyd-aulic systehl, 9-0 rotor brakt., "-3 mold castings, metal, 17-2
Mylar. 2-9 ½
~M-vmanent
Pkint~s, 2.34 to 2-38 Paramewnr, rotor blade tkAisn, 5-37 to 5.41 Passenger compartment, desifin of,. 13-5 Perfor.-arag antitorquc rotor, 3-82 h~licepicr., " system parameters prcejAx%-r
S4-4
I-I
'Piwtt-wiv
OGAtst wa.o 16-32, 16-13 bydtamlc Mrwsinrr, as&."my f1W adobea pump opmrtlin, 9-19 types and uu- af 163., 17-12 LI,' Lrauii tyctum oign. 9M K4A contro6, desi~p of,13-2 wib/ rlatb has.;" design W1 tostift of. 6-IS to oi-Zl1y,21ks~mo9rtq trad,-o. 0-10 subs*.ium. dsign ar., Pr %.x,2-3 is-ýakiition, 9.4' to 9-41
Pitoot-9tatic f .ýbs I-44
Pr&oa at flight puths. £4.3 PrCisile
Yftot tubes. 94i Plaform, Frpmpbr blink 5-611 Plasicscom,,-ats appli4Ltions for, 2--7 diakdvantages c-%, :-7
4
bVAdw 549 to 5-75 See AW Dadev. pvpropur (hubs, actuators co-Attwsk $5.6 to 5.66 dpTAnici, 5.57 to 5-65
~a~-ei~oc~,2-11, 2-12 reinforc -O.deaigr cosideaticxi*, 2-11I to 2-mt Plating *kQzrc!"s, 2-37 delctroiytic, 2-3? zwtal, '-36, 2-37 molten Mets! Jip, :-3S. vs.-mum dvn..iaition. 2.37
fineumatic wuayste~e analys~is of~ mhs ~iow. 9-32 caipeneni desi~,, Criteria. 9.33 to 9-44 des4n re~uiramzu~t, 9-32, 9-33 canos rddhdzor,9-354
-s.
TAV-
--..
'an"
'
.wa
.
Pumpsh-lcnrl Searbox lubricating oil. 4.63 hydraulic, 9-18, 9-19, f-30 advantagia of. 16-28 compontattu of, 16-28
;rhstaIlab..'a and qud~ifica'.oi. 9-4# guainuiina~biity '-nsiderations, 15-7 piolvidt' sydern, 9.45 to 9-" prvuim- rcduo~i~z end Pagulatioi-, 9,38 *,,rnisfi
ykw iPrquiramens 547~ Maintteinabitily coaisideavions, 15-5 ktlafti&4473 RoptiI§.(o sUbz)%UM dceiintiva oý, 3-1 interfac~e with C3SE, 15-3 quck-cbave cqmbifity in, 15-4 PNd~cy Sun,%uns of, 16.27 Pulleys, Conuol system criteria for selectiogi, 16-26 installation of, 16-26 vvpsti and vtas, 16-25 to 16-27
P.ga,2-P
moittw~
N~Y
9-34a~
t -3ting, componci. is awl -nstall',tion, 9-45 tyN. &id claw.r~ 9-.44Q X\yte,2-9, 2-12 I sition lights, 13-10 Pci;rer coiatrols, criteria for, 6-14 Po ver/lt.~ relationship, transmission and drive s3 .Lcem components, 4-25 Power losse, transmisision, due to acces' ie, 4-11 Pc,aai source, electrical. selr'tion of, 7-1 Pr-pisi GRP, 2-11. 2-19 PrM. vre S8a:s. hyd.-au~ic, 9-30 pneumatic, 9.41 Press~are reguktio~i hydraulic sytr.~emw, 9.20 Pneumtatic wy~ww. 9-38 Ilasammr bcnsi& reurun line, hydraulic rsystm, 9-9 F~wsurv, vuausas, pneumatic system, 9-42
type
and.
6-I
161
to
16-30
6~7t163 q-feel system, definition of, 6-14
Qual ' ication API), 3-16 transmission and drive system, Lu29 to 4-32 Quality control propeller blades, 5-73 r-jor Wlade construction, 5.45 Quick-disconnocts APU, 3-16 diesse.47 GSE iInterfaces, 15-4 pneumatic system, 9.a44 Quills. gearbox. 4-67 R Rate controls, description,64 Rating% transmai~on, 4-25 to 4-28 Reduction drivois, APU, 3-2D
waS sd wad pnm babs.a
aukis" afrms rowkioiNs 14-10
wl,
raw.. w inadvat w~k nuiisia4
Mhdwof44w=Fk
*men. #47.
a
O~ubmil, SA.61ONa5u
APe. wiAaly
~~k RVoWor
49 1~
wok541in f 5-78 wiu4.-
Ro-57u
t-510'-i
Bys.~U~ at~ ydradc for Ruuirciors ofotor Vea~igrol, 449i 4-72r oostoof~.4Iso0;
2
at.i 6- 1.6-22
13 msAsile- 14.86yt r~til-;, *iV4gu WAiVkels, 14.
\\YP
fz
RcwpIgs ii
byWA WMAM.ut 9-25
typvai iod dusW Se26 prigidrhockort
wl Iea
hu ewkifmt:c
AIPU. 3-19
!r Rlbwui
-4
4raw6ra
48C6d"l
Rob"
Rvels,
(Ms. 6-7
9*2
or bldeladi f,48 1 2 khw1oio.
-4 dASs
ii
15 F;i-ity9 hixdoa, %nU*iw 5 ta5s 1.p hing
~ito53
Soagn* wameters
Sckl ,to 5-78 4C
2-9 o,212-O&~rcl.ilt~l~l
eaacmpuitc.'
144 @ ilieg@6
wtur9
3542
nmuiv--o
52
, 1382
~ 41
fim~ ehowawimSimm desig
asgglii 4C&VWW
Axwwl for 41d 2elaph, 4ham.5) 22
am6I7-
Sadg.'t-,'-,.~me~~el
S.*lanta mL2-30 23 do!"an. ols 2-36
m643u SaSk dwiopiavnm, - -f -lh -a~o di, .27 4.1
23 xface, 1645%2-3 km hydrauic, 2cuao3,4-I sd,'rinth.2-34 umbiaine, 2 rb34 44 r~~iaib~atio rait-ic aka ofratps n merice 9.25a fcarcit.16-S,
Skodabmiag Wairam12 114 Sokeing. descnripio
hrtvured, Nituk-1 curkzwd. p3-3
Speei44dgoesrnpiog
-1
sysevM, in Ssors, f~kag rain S~crm~13
W
kslmdUS%3 dm4., 32-9 iaglaImmme 6-17 Smn okig p~imoist.-134
amatypes of16.15
984
U,3.3
wtrol systc'm, 9-24 ~~f'2 coni SAS, hydrraulic subny~teni. 9-25 S~batiogdrawings, intersoinnect, 4-73 suibcritical, 4-74tpe,-5 &wapercriucm, z..76 taqt rotor or proptlliw, 4-73 trmsnsic:ion &nd drivye system, 4-81 Shafts, flexi~,e ccairol, IS-30 power, 16 0maneuver, tyL4 characteri.tics, anid ionstrtic*.ion. K-29. 16-30 Shel, ume of, 1-1 Sheet metal 4urmint, rlscription of, 17.3 lftkiding, ekctricsl, 7-23 .Th;-,y, wheel, anijysis of, 1248 Mbock strub, ianding gear, U2-5 to 124 Shot 'maink pSuet~ 4.40 wo k hardOWin of mnAWa surfaces 17.20D
~Sightins sutions, *.apone stim, 134
rfcs,1 1-
Skitulndprotngparead aungm, 12 uptUii 39 3 n 12-10sc~ru Perovsk for 13-4 SM040shaid~ ,ufxs 11. Spee stailt5
Servy"
(I
vmiim S.36,k 547w
Spinning
n or.37-19 requirements, f-1r13-
7,me4a-48mi
1-
Spiersal rol, 6-, 6-6remsfo.6 Splicing, cable. 17-13, 17-16 Splines, power transmission, 4-57 to 4-60 analysis of strength, 4-59 4-60 properties, 4-58 Sprayup, GRP, 2-1S Stability degradation of, 6-8 derivatives, wind tunnel testing for, 6-2, 6.26 inherent, 6-6 criteria for, 6-7 measures of. 6-4 reqaimrnmfl' for, 6-,' rotor blade motions Sm.- Iuiamility toinonal rotof and drive system, 5-26 Stability ausmetntation symtm (SAS)
coat considiratong 6-133
criteria for sulction of. 6-10 design coedaio,6-1 to &.9 sedhe of helicopter aiWs (P-30 OhctrobydnrUMC. 6-1t) fiUre. OiW rpoem 64. 4.9 6-1;.
hw~sianat.64Jsupporft.
~I~S11Iuip 4-I
Sirvwaih hydisialic show ?F'1
6.12 dMu
K4&
e"u M"u
11-2 .97
iad4sodseuhn Lf A-19 ewnbufioa Wo4-17. 4-13 mad.s by. &I3ISutvivalAdmaNPS .qnlminw
~wnSof vhd
imant 9-
~uudeal
Sriva
ftl &OS. Make. Paason tswk a" akin w~iu addM ovsad sad wandisuu%4-29 5twb'. DC, 7-11
p
13-5rog1o,
Oav~aa
Mlmii., 14-2 17-15, 17-16 Swvagi& balok
rotary, 16.22
APnd-12
Mule, :6-20
typ (qkdvieW4 1-11
kAboas
"ima-fr.7-
3.
A) for, 9-2
hqa~mu appklikai to sixtrical sysaw dwign. 7-17 9-3 stnaosug Staic ohecuicity, ProwcwMoaofaekiricalsea sshAks, 7.29. 7-3 1 Static woads cargo co. aJkwastL 137 cuenss uorgo, MIS3 systfam 8-7 SStatioankeapiqt aloy. 2-2 Caro,coviouion resistant (staiakss, maraging, 2-3 prodpittion hardenitq, 2-2 Stiffars 1-6
2-.236
stiffness
airframe structure 11.3 Moigt cootrol symw%~t 6-24, 9-6Terminal Stiffea*4-wesight ratio.,occm artn~u.e 2-21 Stop., rotor t!*de. had-ag. zop n flat, 5-4 Storap boules ponAnltic systern, 9-12 Suina~ner mom~ feed syrnan, 3-13 Stres relif Iin beat resalmoi, 17-17 Stretchfomin~g. meal ýewt 17 3 Umrapsa. urfmraja smarijg,
11-6
Str,.'. sitr-aii dealgn of. 12.3 to 12-4 SUMP f~ aL3-11 FtOI4dfmlq 4-22
APU. 3-21, 3-22 armament iLyasaa. 14-10 blade fokding coasido-mime.a 5-36 flight control, 6-12 T Tactical air navigatior~(TACAN) sYst~ms, &46
Toil rotor shafting, 4-73 Tail rotors &v': Rotor svstcm, antitorque intislip, 2-37 high-visibility. 2-'77 marking, 2-37 types and w3, 2-37, 7-3L Teeterna rotor See: G~mbaled rotor TO'Thn in bearings. K-13J
Tivmpering, 17-11 '7nhiion-torsion stain. rotor blade' retentions,. 5-32 maneuvering equijpmen.. 8-5 Tri~aninal arips or boa:ds, 7.25. '129 16-19 Ttfinals. electrical, 16-19 Tcst rolalts. provpulsichn syrtean. 15-4 Teating developmenit, SAS, 6-13 trarsrnvizisuit and drive system, 4.29 endura'noc liakit, 5-54 to 5-16, 6-22 fuel system., 3-13 )eummiac ryatzm, 9-45 structouull, 31-13 I-is
INDEX (Casmhwui TONS airspeed cahsbrabion. 9.47 cautizn tr~anuna nD drm# & vssy~aa 4-29, 44$ delcto.,gs ~g~a,4-29 sendtarmaw bywdrm c actuator. S-.15 ftigbz load survey, 4-74. 5 Ill, 1-06. 6-22 vasrbox imembly and diww&.bly. 4-M~i ymchax officitaq assiumiweta, A-A, 4-31 hypuanim pu~mp cou~paibility, 9-19 lhbicatima tranmissionmosd drivi. a ybtwtu 4-30 modeL, waWa Lzw4ing, 12-14 7 over~vww, yearbox, 4.32 pmftU~ht a=Wpt4ac (PFA1). of groad NOs vwvle (GTV-), 4-31 thmwnW roappl.g, Searkx, 4-31 J4Tberumoj&uic axticiias. 2-4 Thsnnocctiig rwiia, 2-9 Tic baus Saee Tfvwxoo-torsion strap Tiodown devices cargo. 13-15 1;nCbw w_-owCr4.w'% %MIO).tfanaauiuuor Owd diwv, ryzuri~s 4i-27 Time. landiiw rer 12-3 Titar~~ii~ a~u~ csafaactcistics, 2-6 machie-ing. 2-6 Telerencem. machining, 17-3 Tooling airfrai, . 9aztwti&, 17-1? idejýSn toquimmeroet. 17-20 to 17-23 oputul. 17-22
copaft aiow, 4-23 to 4-25 "h..am cnllgatios for muvivabitiy, 4-1 to 4-Ul sludkfing ww~ atoeCt GIyiLt 4-72
to 4-4:
d~ciwamcy. 4-4 to 4-11 faii-wr modes, 4-32 to 4-34 aemJrquirer.ata, 4-3 to- 4-23 lubrication, 4431 to 4-M noise levhLs 4-11 qualificetion roquicamemdu, 4-29 to 4-32 rating., 4-25 to' 4-26 rviaiability. 4-12 to 4-17 s3., 4.1 surviva~ty.4-17 w 4-23 weiglit 4-4 Tianamitter, pnswruw hvdnudic WyMMn, 9-fl Trausp~aiencas, cockpit 'and cabin., I -I I Trim symans continuousa, 6-26 wi:~ntainabilit'y cousideratiows V5-5 par", 6-26 re~uirwwn=tA and types, 6421 T%6~~i.mhydraiAi~u aili_ 9-26 Turbulene,- ioffe,; on wimbility ax cnd il 6-7 Turnbvcklsýk$ uses of. 16-3 U Urcthanes, 2-9
propcAlcr WAade, 5-72 rotor blade. -45 inpuat (radhnc) lii~tz tranan'uiaon iin drivr. sux4-2.taA-rotor shafting, traruient critraia, 4-74 Toxic gam~, pre'iczowj rron-, wt'?.ona ihiswaiauonb, 14-i L Teansient ra~ponme &scii'tion and damping requiremuents, 6-6, 6-7 msnarcuvcrs, 6.-7 pulv.; inputs 6-7 Trvinrm'ýiniq~1
V.?.vca bWood &nd back pmnum', air com.pressor, 9-35 control, whw.d brake. 9-28 Iiydikulac, chock-, 9-21 control scleto, 9-22 d~rv "-itcrie. (radditional), SP-1 pressure relief, 9-20 mastet control. hydraulic actuator, piacu.matic, 9-36 to 9-41 directional ýor~trol. SOB8 to 9-41 whock brake, 9-44 Ventilation, cockpit sacd cabin, example Rnalysio.
staoartl,, 4-21 thermal ma~pping, 4-31 See also: Oearboxft Trap.;mihsion and drive subrsystem inclu&sd ina, 443, 449
VentL fuel tank, 3-1 1 swat ps.-aure, syz'ean, C-A7 Very high ri~uency omjh eqs (VOR) ) utgas.na V.5 Vibration isolation
dynxmak, 4-34 to 4-64 static. 4-64 to 4-72
A~tgi'ft ins ti~uin, 3-5 instriarient ingallatice, 10-10
Vbnigmmee~auI~.tvm hrna ad Vb,~laasWoo= * ouu~wuo I ar'asw dufta. 11-2 rawI. 5-19 su Mdr " ft m cya ostrsd of. 446 le 4-70 £ndbvms raw, $-a 11-9alm 3-57 ta 349 dar to man md anceftamic W~an 544 rotor, amid"cmicoim 5-16 Vhibfty ftoq ts"S, cOwkp, daifta. 13.2 Vokag.. r~doan of
AC. 7-19 DC. 7-10
UftKY o Md iu
asm, 14-10 I 16-.1
ft Da~ F~ %A,14 Wwu Ptabws hiadhif P. 13 ow-IgJ Cobsil.cI &WtI 6014P pbms 1-I U.umi~mb ind drAm l- a 4-m. 44 441 wow"b rawf mbh two* Sai bum. 3.44, M.'a W.Idh& tpsnm OW~j &%ps 1747 WIILhuu.Inmbg p~ar BAmig nwVnkvim~aus 12-4 kI-%dio, U~-2.12-3 Wbalnb, prouad hIudO'q, 124 WiGGINAM, Car
~dmA fligt coaxrd. 6-13 MS.6-15 dds /dskiciu 13-4 trmamlm addrive syshni 4.17 to 4-23 &W du048a of, 11-11 mcthod of analym. 14-13 to 14-16 vi~o requirsaz.euas. 1342 Wadctioti of. 14-V to 94-20 Wind-e&nu1~ toolit4 6.2. 6-14, 6.26 ch~cki~
for.14-16nad1.s,
wWann War~~iin lishl. hydraulic systm failure, 9422, 10k.2 &igtals. 10-2 trasn~miuaon iwnd drivs sy~su=, 4.70
voi., W02
Watef laninsg cpabihiI darig ad t &vciopmmnz for. 12-12 WtapOn ilns~llai~ols gvided miuskilco 14-6 to 14-4 sung, 14-1 it. 144-z tickcii'. 144 to 14-10
7-25 wwificati-mw 7-22 t.YPK, chAracasTiui.a NEW sppicatozia Wim. ptCrviumvA 16-11 inMkty, 16-32, 16-!i stavcthdMI 16-31 to 16-33
7-27
to moVI4m Compo~nents. 7-26 WOek hardwsin#,ý fflaluý 17-19,1I7-A) Wrow~ht ptodueta. alw'rtaium ally. 2-vZiow piatiml 2-36
1-17
AC
FOR THE COW'MMDER:
I
I
ROBERT
~OFFJCID
G.J AROLD LTC. Cu Adjutant Gmne'-a1 D7STý?IBUTIOM:
S;,eL ialI
wOvme3mmI
il
L.KIRWAN
Brloadier Geneval Chief of Staff
MIS~TIWUOVKI
i--'nx
413
S
O-O
_-----_
_
------ bowf
*
IPM
... W esC j11W" ID.-1 9a ." 'nuns4~W St. 9k ft&. me %..t' WADI"sl tg1.9, e Ctlna . hps l.w . DI . ft""-. 'I est-s.. "mixi Teatsi IA.fl ,wee hwa. Suts eet hC ItMMS I.94, wi.-s4t6 Mmen. .S.. O-..aWd 0 * Usft..S t stlt A.ul
mane .
36
bee
t"M".
m =1 &AD
49
la..lbt W.
St...-
.
II
116
helela
fall-
1hupme-S
Iký e.
hu
sa.a.a.
n..
ll.
all
114
aws..
bme-..qII
.ea-
aýI
I lsl
U lstr_ '.
.*
fIt .. ) eta.. fet.
eww
Sit... ff
.4
,t-'
~~
9
I
'll) Iný
A
ee.()
Sw~e III .
~
1IUa Off
nA ... 1
&asutI..1 am
IV?
q
tt.Pwl..te. iele. vat
u.eI Cs.
Iet.".
1 1t t
et.
V
a-)
I
et. 19.Plt.t. 1MeteShea--M awas" -1 W--tI el sIw
Se~n
:t
dI~t
n e(9
wl 911
~e.s II.hu*Itt
ajwMseS l b Selee
ta." n1.p iaetl.*
1 11.
SI 4
t
It....IS)(SO
w
a-N t-Aai~*
..
DOWn)ael
iW.-t.' G.eeP
M-1-
(an.-i*,4..jlllDP ht *' s
I'C..l.t-tstet
l,,
if~te*iL~l
e. 11a)t b~et .19 flub. lnt S oOtnl.aU..
r.
9.adl:
-''..da k.1.411 I_-
tl
I.:C,.
-- I.1eete aI. s .w,*~.cet~.~-. timere
tI.s.I
~
.-ww9.,
ft4.
vpt. Tht
.. ule1
ft-s =
-11-n u
-e
~
e.~ ae~e. e~ln~tl. .. I.q.e ee e ,wa. ft1e,.
If
-d
.e.tft..s.I~e. e~tAla 4.ew~l huh. 9.t. 'tt fi-.I..l'.Uit.) ctet..._-ttt~t. eSe.t
CAe..eeitAS tr lllMW ken.,' " .. W iw
7,tx
.hulas
a.1.
IN) .. wr-tI..e.
a.- It-s
... ha..C. Pta).. (
.*'
R. 4 a.'99,"4 tf~tl
*9~~.I.14
WI
I.
I"I
Sa)
1L1te
a ms.Add
"Olh
Ffetww'
e.
1lI-41w, PwewHdUI.
w II
SaIfta4 . rer1
.n-.ir eanw.9.
wlw.I.l~n t
...
I.taMil -. .. a U~i..M. wl, I 10(1' 'q S &±54-tl,.4 tuettse
I
'a-,
(01
Use...
I
JIfr
ee
Pa.
w Jos5')
II
...
i..
b-.kenlas
3*9)- ee 4,
ml~ntlt
It
lte
I
hu~i
USi ;v 4'~~~.
MM=.-ta41,
1se.w Thee. a.ll*f
e..
ad
"3' I
tays
a..-
i
hs l
.9 7mtucE -d
fweer
tW ..
*--* lM-
4.ablell.
I.tt
fr,.tlt a ud
. a,.eI ti.1a)71t
as*~
£tl~w
e#9hll
-a
huft--Lmt&a
of Ibwl..94 -i 3WAM 69.. wwel.w . a ... 1~l . lan.,..
p1
Tn
shl....I.as.
1.4
is
ect.
11.-
htal~
in3
Ut*
,
111O-.Me--Istsl,
ozt=..I
,-I
If 7
1
an-114 Inttets fa rh~ln.Ae. " vena~sl I, %,s ul . s, .-t
9115 t*...a.4n:6 b.l. . mans.t sempt .... 6,..
s *4
I.....
ft... .
fw.,~~le If$~11 =
off-
...
huina.S
lam
ht..`L
%.
%.( s
mII
S-tO) .. atls* I.., Asm-tS1.
14tt1) s
ft"
Wflr
. fe.9lwa.. . b .. sIpft- .
1-#e. M)
.l.U.A
a.
P.6-n ?gut u t.%g3u~l.l.wus Gater-. '9 .Calrupt .&W vew.. .t
DISD 1
not.
te#9
ý a.ore-l. C .=1.91.
.9
. ,
(U) 9
nft. fteI.
Ue
5AsP*
.
beM fin69. heat. *9 ft G-.'..hu. = ad ats.wi tI. CItie.s a 151... fA
U
. 11(.. .a., ~~~~~~11301
a.
. ft-
t.b.w .9 huts
C
mu3
e
-A.
alyt.St-M i..
flnt an
esigna' Dt
of
14
S..Ilva(S.. w.,Mta'
.
~.tu
Autometiva
lia
