Y USARTL-TR-7- 2 2 / y

AIRCRAFT CRASH SURVIVAL DESIGN GUIDE I

VOLUME IV - AIRCRAFT SEATS, RESTRAINTS,

LITTERS, AND PADDING

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PPLIED TECHNOLOGY LABORATORY POSITION STATEMENT . -

This revised edition of the Crash Survival Design Guide was

prepared for the Applied Technology Laboratory by Simula, Inc.,

under the terms of Contract DAAJ02-77-C-0021. The original

Crash Survival Design Guide was published in 1967 as USAAVLABS

Technical Report 67-22 and subsequent revisions were published

as USAAVLABS Technical Report 70-22 and USAAMRDL Technical Report 71-22. This current edition consists of a consolidation

of design criteria, concepts, and analytical techniques devel-,.,

oped through research programs sponsored by this Laboratory

over the past 20 years into one report suitable for use as a

designer's guide by aircraft design engineers and other interested personnel.

This document has been coordinated with USAAVRADCOM, the U. S.

Army Safety Center, the U. S. Army Aeromedical Research Laboratory, and several other Government agencies active in aircraft

crashworthiness research and development.

The technical monitors for this program were Messrs. G. T.

Singley III, R. F. Bywaters, W. J. Nolan, and H. W. Holland of

the Safety and Survivability Technical Area, Aeronautical Systems Division, Applied Technology Laboratory.

Comments or suggestions pertaining to this Design Guide will

be welcomed by this Laboratory.

DISCLAIMERS

The findings in this report are not to be construed as an otffc,al Depter.ment of the Army position unless so

designated by other authorized documents.

VMen Government drawings, specifications, or other data are used for any purpose other titan in connection

with a definitely related Government procurement operation, the United States Government thereby incurs no responsibility nor any obligation whatsoever; and the fact that the Government may have formulated, furnished, or in any way supplied the said drawings, specificetions, or other data is not to be regarded by implication of

otherwise as in any manner licensing the holder or any other person or corporation, or conveying any riglhts

perrmision, to manufacture, usll, or sell any patented invention that may in any way be related tnereto.

Trade nomes cited in this repors do not constitute on official endorsement or apoloyal of the use of such

commercial hardware or software.

DISPOSITION INSTRUCTIONS

Destroy this report when no longer needed. Do not return It to the originator.

NOTE: This is a revised edition of the Crash Survival Desil 'S

Guide (formerly USAAMRDL Technical Report 71-22). A•,

previous editions are obsolete and should be destroye. -

.. . ....-. ,-.- .n ,..L.,.,,s.•.•diI1•4 -
SCCURITY CLASSIFICATION OF 1HIS PAGE (Whon Data Eftlm.e0

~'EPORT DOCMENTATION PAGE BEFORE COMPLETING FORM

!Ufj 40 4

-*C E.SMED

/0 S.. P.?ISTO e SjATEMEnT ofgl.RfAppove fr pbli rleae;n disribtio unlimited.021

P. SJPERFO MING TANýAy O NAEADADRS 0 PORMELMNPOJC.TS

Vomlum Incfafievl.erpr

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Approved for pulcrees;/itiutonlmtd ~ ~ 
611CUt1TY CLAWaFICATIONM or THIS PAgg(Ubw Oat* bdh,,

20. (Continued) " Volume I - Design Criteria and Checklists*

Volume II - Aircraft Crash Environment and Human Tolerance-' Volume III- Aircraft Structural Crashworthiness8

Volume IV - Aircraft Seats, Restraints, Litters, and Padding, Volume V - Aircraft Postcrash' Survival.

This volume (Volume IV) contains information on aircraft seats,

litters, personnel restraint systems, and hazards in the occupant's immediate environment. Requirements for design of seats,

litters, and restraint systems are discussed, as well as design principles for meeting these requirements and testing for verification that the systems perform as desired. Energy-absorbing devices for use in seats are described, as are various types of

cushions. Delethalization of cockpit and cabin interiors is discussed, including the use of protective padding and the design of

controls for prevention of injury. Finally, computerized methods

of analysis for evaluation of seats, restraints, and the occupant's

immediate environment are presented.

Acesion For

MC TAB Un~mounced

r)tI

UNCLASSIFIED 6ICUSITY CLASSIFICAIION OF ThIS PAOEf(IVh DA1. 5IMOP.)
PREFACE

This report was prepared for the Safety and Survivability Technical Area of the Applied Technology Laboratory, U. S. Army

Research and Technology Laboratories (AVRADCOM), Fort Eustis,

Virginia, by Simula Inc. under Contract DAAJ02-77-C-0021, initiated in September 1977. The Department of the Army Project

Number is IL162209AH76. This guide is a revision of USAAMRDL

Technical Report 71-22, Crash Survival Design Guide, published

October 1971.

A major portion of the data contained herein was taken from

U. S. Army-sponsored research in aircraft crashworthiness conducted from 1960 to 1979. Acknowledgment is extended to the

U. S. Air Force, the Federal Aviation Administration, NASA,

and the U. S. Navy for their research in crash survival. Appreciation is extended to the following organizations for providing accident case histories leading to the establishment of

the impact conditions in aircraft accidents:

9 U. S. Army Safety Center (USASC), Fort Rucker, Ala- bama.

e Civil Aeronautics Board, Washington, D. C. " U. S. Naval Safety Center, Norfolk, Virginia.

* U. S. Air Force Inspection and Safety Center, Norton

Air Force Base, California.

Additional credit is due the many authors, individual companies, and organizations listed in the bibliographies for their

contributions to the field. The contributions of the following authors to previous ediions of the Crash Survival Design

Guide are most noteworthy:

D. F. Carroll, R. L. Cook, S. P. Desjardins, J. K. Drummond, J. L. Haley, Jr., A. D. Harper, H. G. C. Henneberger,

N. B. Johnson, G. Kourouklis, W. H. Reed, S. H. Robertson,

"L. M. Shaw, Dr. J. W. Turnbow, and L. W. T. Weinberg.

This volume has been prepared by S. P. Desjardins and Dr. D. H.

Laananen of Simula Inc. M. J. Reilly of the Boeing Vertol Company provided recommendations for sections dealing with troop

and gunner seats, and P. A. Rakszawski contributed to the chapter on protective padding.

3

-- ~~- - ----

•-__- ,L - ,, ,I II II !I ' I I
TABLE OF CONTENTS

Section Title Pa.e

PREFACE .

LIST OF ILLUSTRATIONS ................

LIST OF TABLES . .. .. .. .. .. .. .. .. . .. 15

INTRODUCTION. . . . . . . . . . . . . . . . . . . ... 17

CHAPTER 1. BACKGROUND DISCUSSION. .......... . 20

CHAPTER 2. DEFINITIONS. . .......... . . 25

2.1 AIRCRAFT COORDINATE SYSTEMS AND

ATTITUDE PARAMETERS. ........... 25

2.2 ACCELERATION-RELATED TERMS . ...... 26

2.3 VELOCITY-RELATED TERMS . ........ .. 26

2.4 FORCE TERMS. . . . ....... .. .. 27

2.5 DYNAMICS TERMS ............. . . 29

2.6 CRASH SURVIVABILITY TERMS ............. 30

2.7 OCCUPANT-RELATED TERMS ......... 30

2.8 SEATING GEOMETRY . . . . . ...... 33

S2.9 STRUCTURAL TERMS ......... .... 35

t

CHAPTER 3. PRIMARY DESIGN CONSIDERATIONS ...... 37

3.1 INTRODUCTION . . . ....... ... 37

3.2 OPERATIONAL ENVIRONMENT ........ 37

3.2.1 Comfort ..... ............ .... 37

3.2.2 Seat Adjustments ........ ... 38

3.2.3 Vibration Damping.. ............... 39

3.3 CRASH ENVIRONMENT. ......... . 40

3.3.1 Dynamics and Kinematics:. ... 40

3.3.2 Design Conditions and Envelopes . . . . 44

3.3.3 Structural Distortion. ............. 45

3.4 APPLICABILITY OF CRITERIA ........ 46

3.5 ACCEPTANCE CRITERIA ............. .. 47

3.6- SELECTION CRITERIA . . .... 47

CHAPTER 4. DESIGN PRINCIPLES FOR SEATS AND

LITTERS...... . . . . . . . . . . . . 48

4.1 INTRODUCTION . . . . . ....... 48

4.2 LITTERS AND THEIR ORIENTATION . .... 49

4.3 MATERIALS ........... ......... . 49

4.4 STRUCTURAL CONNECTIONS ...... ......... 50

4.4.1 Bolted Connections. . . . ...... 50

||5
TABLE OF CONTENTS (CONTD)

Section Title Page

4.4.2 Riveted Connections . . . . . 51

4.4.3 Welded Connections . . . . . . . . .*. . 51

4.4.4 Seat Attachment...... . . . .. . . 51

4.4.5 Joint Deformation. . . . . . . . . .. . 52

4.5 STRENGTH ..................... .. 62

4.5.1 General. . ............. 62

4.5.2 Limit Analysi Concepts. ....... 65

4.5.3 Large Deformation Analysis . . . . ... 66 "

4.5.4 Strain Concentrations ............ ... 66

4.6 RESTRAINT SYSTEM ANCHORAGE . . . . ... 67

4.7 CRASH ENERGY ABSORPTION. . . . . . ... 67

4.7.1 General. . ...... . .... . . . 67

4.7.2 Principle of Energy Absorption -

Illustration .... ............. . 70

4.7.3 Dynamic Response. .*... .. 82

4.8 COMPUTERIZED METHODS OF ANALYSIS . ... 93

4.8.1 General. . . . . . . . ........... 93

4.8.2 Program SOM-LA . ... ........ 95

4.8.3 Calspan Corporation -'CVS .. . ... 99

4.8.4 PROMETHEUS .... ............. 100

4.8.5 Air Force Head-Spine Model . . . . . . . 102

4.8.6 One-Dimensional Seat/Occupant Models . 103

CHAPTER 5. ENERGY-ABSORBING DEVICES . . . . . . . . 106

5.1 INTRODUCTION .............. ... 106

5.2 WIRE OR STRAP BENDING ............ . .. 109

5.3 INVERSION TUBE ..... ......... . . . 112

5.4 ROLLING TORUS. . . . . . . . . . . ... 114

5.5 CRUSHING HONEYCOMB ........... .. 118

5.6 EXTENSION OF BASXC METAL TUBE OR ---

FLAT STRAP ....... ............ .. 119

5.7 ELONGATION OF BASIC STRANDED CABLE . . . 121

5.8 ROD PULL-THROUGH TUBE. . . . . . . ... 121

5.9 TUBE FLARING . . ............ . 121

5.10 TENSION PULLEY ......... ............. 123

5.11 SELECTION OF AN OPTIMUM LOAD LIMITER . . 123

CHAPTER 6. SEAT CUSHIONS .... .............. ... 125

6.1 INTRODUCTION .... ......... ... 125

6.2 REQUIREMENTS . . . .......... 126

6.3 ENERGY-ABSORBING CUSHIONS. ........ 127

6.4 NET-TYPE CUSHIONS ............... ... 128

6.5 OTHER CUSHIONS ................ .. 128

6.6 HEADRESTS ...... ............. ... 128

S. . . . . . ... - .. w .. ..-

'9-- -,;• ... . --.- '•' • •" q- ,, • -_
TABLE OF CONTENTS (CONTD)

Section Title

CHAPTER 7. DESIGN PRINCIPLES FOR PERSONNEL

RESTRAINT SYSTEMS. . . ...... ........... 129

7.1 INTRODUCTION .... ........ . . ... . . 129

7.2 TYPES OF SYSTEMS . . . . . . . ..... 130

7.2.1 Aircrew Systems ..... ............ ... 130

7.2.2 Troop Systems. ......... .... 133

7.2.3 Crew Chief and Door/Window Gunner

Systems.. .. ..... . ...... 134

7.2.4 Inflatable Systems ...... .137

7.3 GENERAL DESIGN CRITERIA ...... ........ 137

7.3.1 Comfort ................. ....... 137

7.3.2 Emergency Release Requirements .. 139

7.3.3 Lap Belt Anchorage . . ........ 140

7.3.4 Shoulder Harness Anchorage ..... 144

7.3.5 Lap Belt Tiedown Strap Anchorage 145

7.3.6 Adjustment Hardware ........ ........ 146

7.3.7 Location of Adjustment and Release

Hardware . . . . . . . . . . 146

7.3.8 Webbing Width and Thickness

Requirements ..... ............. ... 146

"7.3.9 Hardware Materials. . . . . . ..... 146

7.3.10 Structural Connections . . . . ..... 147

7.4 WEBBING AND ATTACHMENTS ........... ... 148

7.4.1 Properties ............. .. ....... 148

7.4.2 -Width and Thickness Requirements .... 151

7.4.3 Webbing Attachment Methods ....... ... 151

7.4.4 Energy-Absorbing Webbing .......... .. 159

7.5 RESTRAINT SYSTEM HARDWARE .......... .. .. lo

7.5.1 General .................. .... 160

7.5.2 Buckles and Emergency Release. ..... 163

7.5.3 Adjustment Hardware. . . ..... ... . 163

7.5.4 Inertia Reels, Control, and

Installation ............. 164

CHAPTER 8. SEAT STRENGTH AND DEFORMATION -

REQUIREMENTS . 166

8.1 INTRODUCTION ............ 166

8.2 RECOMMENDED OCCUPANT WEIGHTS FOR

SEAT DESIGN. ............... . 166

8.2.1 Crewseats. ................. 167

8.2.2 Troop and Gunner SeatB ........... 168

8.3 STRENGTH AND DEFORMATION . . . . .... 168

8.3.1 Forward Loads. . . . . . ........... 168

8.3.2 Use of Design Curves ............ .. 171 4

7
TABLE OF CONTENTS (CONTD)

Section Title•e Page

8.3.3 Aftward Loads. . . . . . . ...... . 172

8.3.4 Downward Loads .......... ...... .172

8.3.5 Upward Loads . . . . . . ............ .. 178

8.3.6 Lateral Loads and Deformation. .... 178

8.4 OTHER SEATS . . ... 180

8.5 PERSONNEL RESTRAINTWHAýRESS'TESTIN. 180

8.6 STRUCTURAL TEST REQUIREMENTS . . . . 180

8.6.1 Static Test Requirements . . . . . . .*. 181

8.6.2 Dynamic Test Requirements ........ ... 188

8.6.3 Data Acquisition and Reduction . . . . . 193

8.6.4 Seat Component Attachment. . . . . . . . 195

CHAPTER 9. LITTER STRENGTH AND DEFORMATION

REQUIREMENTS .................... 196

9.1 

9.2 

INTRODUCTION 

RECOMMENDED OCCUPANT 

........... WEIGHTS FOR .. .... 196 A

LITTER DESIGN .. .. .. .. .............. 19

9.3 VERTICAL LOADS ...... .... . . . . . 197

9.3.1 Downward Loads ..... .............. . 197

9.3.2 Upward Loads .... ............. 199

9.4 LATERAL AND LONGITUDINAL LOADS .... 199

9.5 LITTER RESTRAINT HARNESS TESTING . . . . 200

9.6 LITTER SYSTEM TEST REQUIREMENTS. . . .. 200

9.6.1 Static Test Requirements.. ........ ... 200

9.6.2 Litter System Dynamic Test

Requirements ........... . . . . 202

CHAPTER 10. DELETHALIZATION OF, COCKPIT AND CABIN

INTERIORS ..... ................ .... 203

10.1 INTRODUCTION ... .............. .... 203

10.2 OCCUPANT STRIKE ENVELOPES ........ ... 203

10.2.1 Full Restraint... ........ 203

10.2.2 Lap Belt-Only Restraint ........ .... 205

10.2.3 Seat Orientation .... ......... .... 206

10.3 ENVIRONMENTAL HAZARDS. .......... .... 206

10.3.1 Primary Hazards .... .............. . 206

10.3.2 Secondary Hazards. ............. , .. . 207

10.3.3 Tertiary Hazards . . . . . . . . . . .. 21C

10.4 HEAD IMPACT HAZARDS .... . ...... .... 21C

10.4.1 Geometry of Probable Head Impact

Surfaces . . . .. . . ..... .......... 21C

10.4.2 Tolerance to Head Impacts. . . . . . 210

10.4.3 Test Procedures ........ .. .. .... 21]

10.5 INSTRUMENT PANEL STRUCTURE PROXIMITY . 21; a. ..... )
TABLE OF CONTENTS (CONTD)

Section Title

10.6 RUDDER PEDAL CONFIGURATION .... . . . 213

10.7 CONTROL COLUMNS. . . .... . 213

10.8 SIGHTING AND VISIONIC SYSTEMS. .... 214 S10.9 ENERGY-ABSORBING REQUIREMENTS FOR '

COCKPIT AND CABIN INTERIORS ....... 219 I" 10.9.1 General. .219

10.9.2. Types of Padding Ma~e~i;l; ;n• ..

Properties . . . . . . . . . . . . . . . 222 •10.9.3 Standard Test Methods• . . " 227"

0.9.14 Research on Materials for EnergyAbsorbing Applications . ....... .. 230

10.9.5 Application of Padding Material. . . . . 243

10.9.6 Ductile Materials . .......... 244

REFERENCES .. ................. .. 246

BIBLIOGRAPHY. . . ......... ........... 259

INDEX . . . . . . . . . . . .............. 264

LIST OF ILLUSTRATIONS

Figure

1 Aircraft coordinates and attitude

directions . . . ............... 25

2 Typical aircraft floor acceleration

pulse. .. . . . . . . . . ....... 28

3 Terminology for directions of forces

on the body ................... 31

4 Seating geometry ... .............. 34

S5 Sketch illustrating buckling or

"dishing" formation. . . ......... 53

"6 Static test floor warpage requirement

to ensure seat retention in crashes. .. 55

7 Concepts for release of floordistortion-induced moments . . . . . .. 56

._-,9
LIS O ILLUSTRTIONS JCONTDI

8 Aft seat leg casting attachment. . . . . 57

9 Aft seat leg casting attachment

modification . .*. . . . . . . . . . . . 57

30 Torsional release of joints. . . . . . . 58

11 Fully released joint . . . . . . . . .. 59

12 Bulkhead in-plane warping. . . . . . . . 60

13 Universal release of a joint . . . . . . 61

14 Pin joint releases oriented to allow

rotation around an aircraft roll 62

axis . .... . . .. . . . . . . . . .

15 Comparison of analysis methods for

simple beams ................. ..... 63

16 Seat leg anchorage to floor track ... 65

17 Deceleration-time, velocity-time, and

distance-time curves used in analysis

of seat/occupant displacement with

respect to the airframe ..... ......... 77

18 Airframe velocity-time curve ...... .. 79

19 Deceleration versus time for various

components of seat and occupant. ..... 84

20 Typical seat pan, dummy chest, and

dummy pelvis response to vertical

crash loading ........... .. . .... 86

21 Initial condition, no load ....... ... 88

22 Onset of deceleration load wherein

pelvic area is responding to deceleration load, but the upper torso

and head are not ............... 88

23 High-deceleration load, springs

compressed ........ ............... 88

10
LIST OF ILLUSTRATIONS (CONTD)

*Figure Eno24 Duration and magnitude of headward

acceleration endured by various

subjects . . . . . . . . . . . . . . . . 91

25 Twelve-segment occupant model. . . . . . 97

26 Occupant model contact surfaces ..... 98

27 SOM-LA energy-absorbiing seat model. . . 98

28 CVS graphics display model . . . . . . . 100

29 PROMETHEUS occupant model. . . . . . .. 101

30 Three-dimensional head-spine model . . . 103

31 Lumped-parameter model of seat, seat

cushion, and occupant. . . . . . . . . . 105

32 Crashworthy troop seat . . . . . . . . . 110

33 Troop seat tension energy absorber

including characteristics for two

wire diameters ............. 11i

34 Tubular strut wire-bending energy

absorber with force-deflection

curves ..... . . . . . . . . . . 113

35 Inversion tube concept with typical

force-deflection characteristic ..... 115

36 Comparison of dynamic and static loaddeformation curves for inversion

IS ~ ~tubes . . . . . . . . . . . . . . *.. . . 116

37 - Rolling -torus energy absorber. . . . . . 117

38 Illustration of corrugated aluminum

foil formed into annular column. . . . . 119

39 Comparison of dynamic and static loadelongation curves for stainless steel

tubes . . . . . . . . . . . . . . . . . . 120
LIST OF ILLUSTRATIONS (COND)

Figure page

40 Comparison of dynamic and static loaddeformation curves for compression

tubes. . . . . . . . . . . ........... 122

41 Illustration of fragmentation and

rolling processes in tube-flaring

device . . ..... . . ............. . . 123

42 Tension-pulley load limiter. . . . . . . 124

43 Basic aircrew restraint system . . . .. 131

44 Aircrew restraint system, including

reflected shoulder straps. . . . . . .. 132

45 Aircraft troop/passenger restraint

systems. . . . . . . ........... ...... 135

46 Gunner restraint system. .......... 136

47 Inflatable body and head restraint . . 138

48 Buckle fitting attachment and motion

angles . . . . . . . . . . . . . . . . . 140

49 Pelvic rotation and submarining caused

by high longitudinal forces combined

with moderate vertical forces . .. . . . 142

50 Lap belt anchorage geometry. . . . . . . 143

51 Shoulder harness anchorage geometry. . . 145

52 Load elongation characteristics for

MIL-W-25361 (Type II) polyester

webbing for static and rapid loading rates . . . . . . . . . . . . . . . . . . 15.1,

53 Stitch pattern and cord size . . . . . . 153

54 Stitch patterns tested . . . . .... 155

55 Wrap radius for webbing joints . . . . . 158

56 Webbing fold at metal hardware

attachment ...... . .... .... 159

12
j• LIST OF ILLUSTRATIONS (CONTD)

Figure Eae

57 Aircrew restraint system . . . . . . . . 161

58 Reflected shoulder strap restraint

system . . . . . . . .. . . . 162

59 Seat forward load and deflection

requirements for all types of Army

aircraft (forward design pulse). . . . . 170

60 Lateral seat load and deformation

requirements for all types of Army

aircraft ........ 179

61 Suggested method of applying floor

warping for static testing of seats. . . 184

62 Suggested method of applying bulkhead

warping for static testing of seats. . . 185

63 Static load application point and

critical body block pelvis geometry. . . 186

64 Requirements of dynamic tests if

Ssubstituted for static tests . . . . .*. 189

65 Dynamic test requirements for

qualification and for research/

development testing. . . . . . . . . . . 190

66 Example of input pulse for seats having

less than 12 in. of stroke . . . . . . . 192

67 Graphic approximation example. . . . . . 194

68 Litter downward load and deflection

requirements . 198

69 Litter forward or lateral load and

deflection requirements for all

types of Army aircraft . . .. .. . . . 199

70 Full-restraint extremity strike

envelope - side view . . . . ..... 204

71 Full-restraint extremity strike

envelope -top view ..... ..... 205

S .. 13
LIST OF ILLUSTRATIONS (CONTD)

Figure

72 Pull-restraint extremity strike

envelope - front view. . . . .. .# . o 206

73 Lap belt-only extremity strike

envelope - side view. . . . . . . ... 207

74 Lap belt-only extremity strike

envelope - top view ........... 208

75 Lap belt-only extremity strike

envelope - front view. . . . . . . . . . 209

76 Wayne State Tolerance Curve for the

human brain in forehead impacts against

plane, unyielding surfaces . . . . . . . 211

77 Measured head velocities in sled tests

with anthropomorphic dummies and

cadavers ................ 212

78 Antitorque, or rudder, pedal geometry

to prevent entrapment of feet. . . . . . 214

79 Pilot/gunner station occupant strike

in-board profile 4-G impact. . . . . . . 215

so Crushable eyepiece concept ....... 218

81 Impact behavior (headform deceleration

versus speed) of three padding

materials. . . . . . . . . . . . . . . . 220

82 Impact behavior (average dynamic _

stress versus speed) of three padding

materials. .

83 Minimum tensile strength versus product

density for polyethylene sheet and

plank . . . . . . . . . . . . 0 a 0 * , 223

84 Recommended stress-strain properties

for padding material for head contact,

with cushion thickness up to 1.5 in. . . 230

85 Stress-strain curves for polyethylene

foam • 232

14
LIST OF ZLLUSTA TZONS ICONTD)

86 Effect of density on stress-strain

curves for polyurethane-foamed plastic. . . . . . . . . . . . . . . . 233

87 Evaluation criteria for energy- absorbing material . 234 "88 Load-distributing material evaluation

criteria . . . . . . . . . . . . . . . . 236

89 Response of composite with urethane, and polystyrene foam, compared . . . . . 238

90 Typical response of plastic foam to compression test ............ 239 91 Specific energy absorbed to maximum

strain versus crosshead velocity . . . . 242

92 Dynamic crushing pressure versus impact velocity for tests at two temperatures .............. 244

LIST OF TABLES

Table

I Crashworthiness criteria for the preliminary design process . . . . . . 24

2 Summary of design conditions for

rotary- and light fixed-wing aircraft . . . . . . . . . . . . . . . . 44

3 Comparison rof load-limiting devices for lO00-to-4000Olbloads. . . . . . . . . . 108

4 Restraint harness components load- elongation design and test require- ments. . . . . . . . . . . . . . . . . . 149 5 Restraint webbing characteristics . . . 150

6 Minimum webbing width requirements . . . 152

7.4
*U

LIST OF TABLES (CONTD) .

Table e

7 Breaking strength of stitch patterns

(test series one) . . . . . . . . . .. 156

8 Breaking strength of stitch patterns (test series two). . . . . . . . . . . . 157

9 Typical aviator weights .... ........ 167 *

10 Troop and gunner weights . . . .... 169

11 Seat design and static test

requirements .................. .... 182

12 Litter systems static test

requirements .............. 201

13 Potential optical relay tube crash

hazards ................ . . . . . . 216 I

14 Energy-absorbing plastic foams and

some typical applications .......... ... 224

15 Properties of selected flexible cellular polymers ............ . . . . 226

16 Summary of ASTM test methods and specifications for flexible cellular I

plastics . . . . . . . . . . . . . . . . 228

17 Material summary ... ............ .... 240

18 Test results summary .......... 241

16 *
p INTRODUCTION

For many years, emphasis in aircraft accident investigation

was placed on determining the cause of the accident. Very

little effort was expended on the crash survival aspects of

aviation safety. However, it became apparent through detailed

studies of accident investigation reports that significant improvements in crash survival could be made if consideration

were given in the initial aircraft design to the following factors that influence survivability:

S•1. Crashworthiness of Aircraft Structure - The ability

of the aircraft structure to maintain living space

for occupants throughout a crash.

2. Tiedown Chain Strength - The strength of the linkage

preventing occupant, cargo, or equipment from breaking free and becoming missiles during a crash sequence. Z

3. Occupant Acceleration Environment - The intensity

and duration of accelerations experienced by occupants (with tiedown assumed intact) during a crash.

4. Occupant Environment Hazards - Barriers, projections,

and loose equipment in the immediate vicinity of the

occupant that may cause contact injuries.

5. Postcrash Hazards - The threat to occupant survival

posed by fire, drowning, exposure, etc., following

the impact sequence.

Early in 1960, the U. S. Army Transportation Research Command*

initiated a long-range program to study all aspects of aircraft

safety and survivability. Through a series of contracts with

the Aviation Safety Engineering and Research Division (AvSER)

of the Flight Safety Foundation, the problems associated with

occupant survival in aircraft crashes were studied to determine specific relationships among crash forces, structural

failures, crash fires, and injuries. A series of reports coyering this effort was prepared and distributed by the U. S.

Army, beginning in 1960. In October 1965, a special project

initiated by the U. S. Army consolidated the design criteria

presented in these reports into one technical document suitable for use as a designer's guide by aircraft design engineers

'Now the Applied Technology Laboratory, Research and Technology

Laboratories of tht. U. S. Army Aviation Research and Development Command (AVRADCOM).

17
and other interested personnel. The document was to be a summary of the current state of the art in crash survival design,

using not only data generated under Army contracts, but also

information collected from other agencies and organizations.

The Crash Survival Design Guide, first published in 1967, realized this goal.

Since its initial publication, the Design Guide has been revised several times to incorporate the results of continuing

research in crashworthiness technology. The last revision,

published in 1971, was the basis for the criteria contained in

the Army's military standard dealing with aircraft craahworthiness, MIL-STD-1290(AV), "Light Fixed- and Rotary-Wing Aircraft

Crashworthiness" (Reference 1). This current revision, the

"fourth, contains the most comprehensive treatment of all aspects of aircraft crash survival now documented. It can be

used as a general text to establish a basic understanding of

the crash environment and the techniques that can be employed

to improve chances for survival. It also contains design criteria and checklists on many aspects of crash survival and thus

can be used as a source of design requirements.

The current edition of the Crash Survival Design Guide is published in five volumes. Volume titles and general subjects

included in each volume are as follows:

Volume I - Design Criteria and Checklists

i Pertinent criteria extracted from Volumes II through V,

presented in the same order in which they appear in those

volumes.

Volume II - Aircraft Crash Environment and Human Tolerance

Crash environment, human tolerance to impact, military

anthropometric data, occupant environment, test dummies.

Volume III - Aircraft Structural Crashworthiness

Crash load estimation, structural response, fuselage and

landing gear requirements, rotor-- requirements, ancillary

equipment, cargo restraints, structural modeling.

1. Military Standard, MIL-STD-1290(AV), LIGHT FIXED- AND

ROTARY-WING AIRCRAFT CRASHWORTHINESS( Department of Defense, Washington, D. C., 25 January 1974.

18
Volume IV - Aircraft Seats, Restraints, Litters, and Padding

Operational and crash environment, energy absorption, seat

design, litter requirements, restraint system design,

occupant/restraint system/seat modeling, delethalization

of cockpit and cabin interiors.

Volume V - Aircraft Postcrash Survival

Postcrash fire, ditching, emergency escape, crash locator

beacons, retrieval of accident information.

This volume (Volume IV) contains information on aircraft seats,

litters, personnel restraint systems, and hazards in the occupant's immediate environment. Following a general discussion

of aircraft crashworthiness in Chapter 1, a number of terms

commonly used in discussing the crash environment, seats, and

occupant protection are defined in Chapter 2. Chapter 3 presents design considerations for aircraft seats, and Chapter 4,

principles for crashworthy seac design. Energy absorption is

discussed in Chapter 5. Principles for cushion and restraint

system design are presented in Chapters 6 and 7, and strength and deformation requirements for seats and litters are stated

in Chapters 8 and 9, respectively. Cockpit delethalization,

including protective padding, is discussed in Chapter 10.

19

19
1. BACKGROUND DISCUSSION

The overall objective of designing for crashworthiness is to

eliminate unnecessary injuries and fatalities in relatively

mild impacts. A crashworthy aircraft alfo reduces aircraft

crash impact damage. By minimizing personnel and material

losses due to crash impact, crashworthiness conserves resources, is a positive morale factor, and improves the combat

effectiveness of the fleet. Results from analyses and research

during the past several years have shown that the relatively

small cost in dollars and weight of including crashworthy features is *a wise investment (References 2 through 13). Consequently, new generation aircraft are being procured to stringent, yet practical requirements for crashworthiness.

To provide as much occupant protection as possible, a systems

approach to crashworthiness must be followed. Every available

subsystem must be considered in order to maximize the protec-.

tion afforded to vehicle occupants. When an aircraft impacts

2. ENGINEERING ANALYSIS OF CRASH INJURY IN ARMY OH-58 AIRCRAFT, USASC Technical Report, U. S. Army Safety Center,

Fort Rucker, Alabama, to be published.

3. ENGINEERING ANALYSIS OF CRASH INJURY IN ARMY CH-47 AIRCRAFT, USAAAVS Technical Report 78-4, U. S. Army Agency

for Aviation Safety, Fort Rucker, Alabama, June 1978.-

4. ENGINEERING ANALYSIS OF CRASH INJURY IN ARMY AH-l AIRCRAFT, USAAAVS Technical Report 78-3, U. S. Army Agency for Aviation Safety, Fort Rucker, Alabama, March 1978.

5. Carnell, B. L., CRASHWORTHINESS DESIGN FEATURES FOR ADVANCED UTILITY HELICOPTERS, in Aircraft Crashworthiness,

K. Saczalski, et al., eds., University Press of Virginia,

Charlottesville, Virginia, 1975, pp. 51-64.

6. Bainbridge, M. E., Reilly, M. J., and Gonsalves, J. E.,

CRASHWORTHINESS OF THE BOEING VERTOL UTTAS, in Aircraft

Crashworthiness, K. Saczalski, et al., eds., University

Press of Virginia, Charlottesville, Virginia, 1975,

pp. 65-82.

7. Rich, M. J., INVESTIGATION OF ADVANCED HELICOPTER STRUCTURAL DESIGNS, Volume I, ADVANCED STRUCTURAL COMPONENT DESIGN CONCEPT STUDY, Sikorsky Aircraft, Division of United

Technology Corporation; USAAMRDL Technical Report 75-59A,

Eustis Directorate, U. S. Army Air Mobility Research and

Development Laboratory, Fort Eustis, Virginia, May 1976,

AD A026246.

20

4-• r l . . i i" .. f ' i "•" -- -"• "•FP -"
the ground, deformation of the ground absorbs some energy. "This is an uncontrolled variable since the quality of the impacted surface usually cannct be selected by the pilot. If

the aircraft lands on an appropriate surface in an appropriate attitude, the landing gear can be used to absorb a significant

amount of the impact energy. After stroking of the gear, crushing of the fuselage contributes to the total energy-absorption

process. The fuselage must also maintain a protective shell

around the occupant, so the crushing must take place outside

the protective shell. The functions of the seat and restraint

system are to restrain the occupant within the protective shell during the crash sequence and to provide additional energyabsorbing stroke to further reduce occupant decelerative loading to within human tolerance limits. The structure and components immediately surrounding the occupant must also be considered. Weapon sights, cyclic controls, glare shields, instrument panels, armor panels, and aircraft structure must

be delethalized if they lie within the strike envelope of the

occupant.

8. Hoffstedt, D. J., and Swatton, S., ADVANCED HELICOPTER

STRUCTURAL DESIGN INVESTIGATION, The Boeing Vertol Company; USAAMRDL Technical Report 75-56A, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, March 1976, AD A024662.

9. Hicks, J. E., AN ANALYSIS OF LIFECYCLE ACCIDENT COSTS FOR

THE ADVANCED SCOUT HELICOPTER, U. S. Army Agency for Avia- tion Safety, Fort Rucker, Alabama, January 1977.

10. McDermott, J. M., and Vega, E., THE EFFECTS OF LATEST

MILITARY CRITERIA ON THE STRUCTURAL WEIGHT OF THE HUGHES ADVANCED ATTACK HELICOPTER YAH-64, Journal of the American Helicopter Society, Vol. 23, No. 4, October 1978, pp. 2-9.

11. Haley, J. L., Jr., CRASHWORTHINESS VERSUS COST: A STUDY

OF ARMY ROTARY WING AIRCRAFT ACCIDENTS IN PERIOD JANUARY

1970 THROUGH DECEMBER 1971, paper presented at the Aircraft Crashworthiness Symposium, University of Cincinnati, Cincinnati, Ohio, October 1975.

12. Hicks, J. E., ECONOMIC BENEFITS OF UTILITY AIRCRAFT CRASHWORTHINESS, USAAAVS Technical Report 76-2, U. S. Army Agency for Aviation Safety, Fort Rucker, Alabama, July

1976.

13. THE ECONOMIC BENEFITS OF CRASHWORTHINESS AND FLIGHT SAFETY

DESIGN FEATURES IN ATTACK HELICOPTERS, USAAAVS Technical

Report 77-2, U. S. Army Agency for Aviaticn Safety, Fort

Rucker, Alabama, June 1977.

1 21
Ideally, it would seem most efficient to simply specify human

tolerance requirements and an array of vehicle crash impact

conditions and then develop the helicopter as a crashworthy

system with a mixture of those crashworthy features that are

most efficient for the particular helicopter being designed.

Unfortunately, the validated structural and/or human tolerance

analytical techniques needed to perform and evaluate such 

a

maximum design freedom approach to achieving crashworthiness

are not available. Furthermore, testing complete aircraft sufficiently early in the development cycle to permit evaluation

of system concepts in time to permit design changes based on

the test results is not practical. The systems approach dictates that the designer consider probable crash conditions 

I

wherein all subsystems cannot perform their desired functions;

for example, an impact situation in which the landing gear cannot absorb its share of the impact crash energy because of aircraft attitude at impact. Therefore, to achieve the overall

goal, minimum levels of crash protection are recommended for

the various individual subsystems. A balance must be struck 

I

* between the two extremes of: (1) defining necessary performance on a component level only, and (2) requiring that the aircraft system be designed for an array of impact conditions 

4

with no component design and test criteria.

Current helicopter crashworthiness criteria require that a new

aircraft be designed as a system to meet the vehicle impact

design conditions recommended in Volume II; however, minimum

criteria are also specified for a few crash critical components.

For example, strengths and minimum crash energy-absorption requirements for seats and restraint systems are specified. All

strength requirements presented in this volume are based on the

crash environments described in Volume II. Testing requirements are based on ensuring compliance with strength and deformation requirements. Mandatory minimum crashworthiness design

criteria for U. S. Army light fixed- and rotary-wing aircraft

are stated in MIL-STD-1290(AV) (Reference I). All pilot, copilot, observer, and student seats in either rotary- or light

fixed-wing aircraft should conf-orm to the re-quirements of

MIL-S-58095(AV) (Reference 14).

Although much higher levels of crashworthiness can be achieved

in completely new aircraft designs, the crashworthiness of

existing aircraft can be significantly improved through retrofitting these aircraft with crashworthy components adhering to

the design principles of this design guide. This can even be

1'4. Military Specification, MIL-S-58095(AV), SEAT SYSTEM:

CRASHWORTHY, NON-EJECTION, AIRCREW, GENERAL SPECIFICATION

FOR, Department of Defense, Washington, D. C., 27 August

1971.

22
achieved while expanding the combat effectiveness of the air- * craft. Examples of this are the successful program to retrofit all U. S. Army helicopters with crashworthy fuel systems

(Reference 15), and the U. S. Navy program to retrofit the

CH-46 with crashworthy armored crewseats (Reference 16).

In an initial assessment, the definition of an adequate crashworthy structure may appear to be a relatively simple matter.

In fact, many influencing parameters must be considered before

an optimum design can be finalized. A complete systems approach must be employed to include all influencing parameters concerned with the design, manufacture, overall performance,

and economic restraint on the aircraft in meeting mission requirements. Trade-offs between the affecting parameters must

be made in order to arrive at a final design that most closely

meets the customer's specified requirements. It must be remembered that for each type cf aircraft, different emphasis

will be placed in the parameter mix. Table 1 summarizes major

crashworthiness criteria that must be considered during the

preliminary design definition phase.

15. Cook, R. L., and Goebel, D. E., EVALUATION OF THE UH-lD/H

HELICOPTER CRASHWORTHY FUEL SYSTEM IN A CRASH ENVIRONMENT,

Dynamic Sciencer,. Division of Marshall Industries; USAAMRDL

Technical Report 71-47, U. S. Army Air Mobility Research

and Development Laboratory, Fort Eustis, Virginia, November 1971.

16. Domzalski, L. P., et al., U. S. NAVY DEVELOPMENTS IN CRASHWORTHY SEAIING, Naval Air Development Center; Proceedings

1978 SAFE Symposium, Survival and Flight Equipment Association, Canoga Park, California, October 1978.

23

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24
4-

S2. DEFINITIONS

2.1 AIRCRAFT COORDINATE SYSTEMS AND ATTITUDE PARAMETERS

Positive directions for velocity, acceleration, and force components and for pitch, roll, and yaw are illustrated in Fig-

-: ure 1. When referring to an aircraft in any flight attitude, it is standard practice to use a basic set of orthogonal axes

as shown in Figure 1, with x, y, and z referring to the longi- tudinal, lateral, and vertical directions, respectively.

+z

Yaw

Pitch •

Roll

Figure 1. Aircraft coordinates and attitude directions.

25

• -. •. . .. . . ... • "• -1j 
2.2 ACCELERATION-RELATED TERMS

* Acceleration

The rate of change of velocity. An acceleration is

in magnitude or in direction. Acceleration may produce either an increase or a decrease in velocity.I

There are two basic types of acceleration: linear,

which changes translational velocity, and angular (or

rotational), which changes angular (or rotational)

velocity. With respect to the crash environment, unless otherwise specified, all acceleration values are

those at a point approximately at the center of the

floor of the fuselage.

* Deceleration

Acceleration which produces a decrease in velocity.

* Abrupt Accelerations

Accelerations of short duration primarily associated

with crash impacts, ejection seat shocks, capsule

impacts, etc. One second is generally accepted as

the dividing point between abrupt and prolonged accelerations. Within the extremely short duration

range of abrupt accelerations (0.2 sec and below),

the effects on the human body are limited to mechanical overloading (skeletal and soft tissue stresses),

there being insufficient time for functional disturSbances due to fluid shifts.

•The Term G

The ratio of a particular acceleration to the acceleration due ýo gravitational attraction at sea level

(32.2 ft/sec ). in accordance with common practice, f

this report will refer to accelerations measured - in G. To illustrate, it is customarily understood

that 5 G represents an acceleration of 5 x 32.2, or

161 ft/sec

2.3 VELOCITY-RELATED TERMS

e Velocity Change in Major Impact (Av)

The decrease in velocity of the airframe during the

major impact, expressed in feet per second. The

.ajor impact is the one in which the highest forces

26
are incurred, not necessarily the initial impact. For the acceleration pulse shown in Figure 2, the

major impact should be considered ended at time t2 . Elastic recovery in the structure will tend to reverse the direction of aircraft velocity prior to

t Should the velocity actually reverse, its direction must be considered in computing the velocity

change. For example, an aircraft impacting downward

with a vertical velocity component of 30 ft/sec and

rebounding with an upward component of 5 ft/sec

should be considered to experience a velocity change

Av = 30- (-5) = 35 ft/sec

during the major impact. The velocity change during impact is further explained in Section 7.2 of Volume III.

* Longitudinal Velocity Change

The decrease in velocity during the major impact measured along the longitudinal (roll) axis of the aircraft. The velocity may or may not reach zero during the major impact. For example, an aircraft impacting the ground at a forward velocity of 100 ft/sec and

slowing to 35 ft/sec before rebounding into the air

would experience a longitudinal velocity change of

65 ft/sec during this impact.

* Vertical Velocity Change

The decrease in velocity during the major impact measured along the vertical (yaw) axis of the aircraft.

The vertical velocity generally reaches zero during the major impact.

a Lateral Velocity Change -

The decrease in velocity during the major impact measured along the lateral (pitch) axis of the aircraft.

2.4 FORCE TERMS

* Load Factor

A crash force can be expressed as a multiple of the

weight of an object being accelerated. A load factor, when multiplied by a weight, produces a force

27
Peak

I

Averaget --

4.0

Figure 2. Typical aircraft floor acceleration pulse.

-which can be used to establisý_ static strength (see

Static Strength). Load factor is expressed in uni~ts

of G.

9 Forward Loado

Loading in a direction toward the nose of the air- craft, parallel to the aircraft longituinal Iroll) ji

28

. . _._ a'_ ... ...
9 Aftward Load

Loading in a direction toward the tail of the aircraft, parallel to the aircraft longitudinal (roll)

axis.

* Downward Load

Loading in a downward direction parallel to the vertical (yaw) axis of the aircraft.

Upward Load

Loading in an upward direction parallel to the vertical (yaw) axis of the aircraft.

* Lateral Load

Loading in a direction parallel to the lateral (pitch)

axis of the aircraft.

* Combined Load

Loading consisting of components in snore than one of

the directions described in Section 2.2.

2.5 DYNAMICS TERMS

* Rebound

Rapid return toward the original position upon release or rapid reduction of the deforming load, uoually associated with elastic deformation.

*Dynamc Overshoot

The amplification of decelerative force on cargo or

personnel above the input decelerative force (ratio

of output to input). This amplification is a result

of the dynamic response of the system.

e Transmissibility

The amplification of a steady state vibrational input

amplitude (ratio of output to input). Transmissibilities maximize at resonant frequencies and may produce motion and acceleration amplification similar to

dynamic overshoot.

29

I

* - - ---. - . ~ - -* - - - - - -
2.6 CRASH SURVIVABILITY TERMS

* Survivable Accident

An accident in which the forces transmitted to the

occupant through the seat and restraint system do not

exceed the limits of human tolerance to abrupt accelerations and in which the structure in the occupant's immediate environment remains substantially intact,

to the extent that a livable volume is provided for

the occupants throughout the crash sequence.

* Survival Envelope

The range of impact conditions--including magnitude and direction of pulses and duration of forces occur- ring in an aircraft accident--wherein the occupiable

area of the aircraft remains substantially intact,

both during and following the impact, and the forces

transmitted to the occupants do not exceed the limits

of human tolerance when current state-of-the-art restraint systems are used.

It should be noted that, where the occupiable volume

is altered appreciably through elastic deformation

during the impact phase, survivable conditions may

not have existed in an accident that, from postcrash

inspection, outwardly appeared to be survivable.

2.7 OCCUPANT-RELATED TERMS

a Human Body Coordinates

In order to minimize the confusion sometimes created

by the terminology used to describe the directions

of forces applied to the body, a group of NATO scientists compiled the accelerative terminology table of

equivalents shown in Figure 3 (Reference 17). Terminology used throughout this guide is compatible with -

the NATO terms as illustrated.-

17. Gell, C. F., TABLE OF EQUIVALENTS FOR ACCELERATION TERMINOLOGY, Aerospace Medicine, Vol. 32, No. 12, December

1961, pp. 1109-I111.

30

I.-' ____ *-- ____ - ____ -. ___.
Headward

(+G Z) Direction of

accelerative force

Vertical

Back to chest Headward - Eyeballs-down

(sternumward) Lateral rightEyeballs-up

Later (+Gh) Transverse

y Lateral right - Eyeballsleft

Lateral left - Eyeballsright

Back to chest - Eyeballsin

Lateral left Chest Chest to back - Eyeballs-

(-G to back out

y (spineward) Note:

Tailward (-Gx The accelerative force on

(-G the body acts in the same direction as the arrows.

Figure 3. Terminology for directions of forces on

the body. (From Reference 17)

e Anthropomorphic Dummy

A device designed and fabricated to represent not

only the appearance of humans but also the mass distribution, joint locations, motions, geometrical similarities such as flesh thickness and load/deflection

properties, and relevant skeletal configurations such

as iliac crests, ischial tuberosities, rib cages,

etc. Attempts are also made to simulate human response of major bcructural assemblages such as - - thorax, spinal column, neck, etc. - The dummy is

strapped into seats or litters and used to bimulate

a human occupant in dynamic tests.

* Human Tolerance

For the purposes of this document, human tolerance

is defined as a selected array of parameters that

describe a condition of decelerative loading for

which it is b,!lieved there is a reasonable probability for survival without major injury. As used in

S'31
this volume, designing for the limits of human tolerance refers to providing design features that will

maintain these conditions at or below their tolerable

levels to enable the occupant to survive the given

crash environment.

Obviously, the tolerance of the human body to crash

environments is a function of many variables including the unique characteristics of each person as well

as the loading variables. The loads applied to the

body include decelerative loads imposed by seats and

restraint systems as well as localized forces due to

impact with surrounding structures. Tolerable magnitudes of the decelerative loads depend on the direction of the load, the orientation of the body, and

the means of applying the load. For example, the

critical nature of loads parallel to the occupant's

spine manifests itself in any of a number of types of

spinal fractures, but typically the fracture is an

anterior wedge, or compressive failure of the front

surface of a vertebra. Forces perpendicular to the

occupant's spine can produce spinal fracture through

shear failures or from hyperflexion resulting, for

example, from jackknife bending over a lap belt-only

restraint. The lap belt might inflict injuries to

the internal organs if it is not retained on the pelvic girdle but is allowed to exert its force above

the iliac crests in the soft stomach region. Excessive rotational or linear acceleration of the head

can produce concussion. Further, skull fracture can

result from head impact with surrounding structure.

Therefore, tolerance is a function of the method of

occupant restraint as well as the characteristics of

the specific occupant. Refer to Chapter 4 of Volume 1I for a more detailed discussion of human tolerance.

9 Submarining

Rotation of the hips under and about the lap belt as r

a result of a forward inertial load exerted by deceleration of the thighs and lower legs, accompanied by

lap belt slippage up and over the iliac crests. Lap

belt slippage up and over the iliac crests can be a

direct result of the upward loading of the shoulder

harness straps at the center of the lap belt.

* Effective Weight

The portion of occupant weight supported by the seat

with the occupant seated in a normal flight position.

22

1-*- _____ ___ ___ __
This is considered to be 80 percent of the occupant

weight since the weight of the feet, lower legs, and

part of the thighs is carried directly by the floor

through the feet.

9 Iliac Crest Bone

The upper, anterior portion of the pelvic (hip) bone.

These "inverted saddlem bones are spaced laterally

about 1 ft apart; the lower abdomen rests between these crest bones.

. Lap Belt Tiedown Strap (also Negative-G Strap, r Crotch Strap)

Strap used to prevent the tensile force in shoulder

straps from pulling the lap belt up when the restrained subject is exposed to -G (eyeballs-out)

acceleration.

2.8 SEATING GEOMETRY (SEE FIGURE 4 FROM REFERENCE 18)

e Design Eye Position

A reference datum point based on the eye location

that permits the specified vision envelope required

by MIL-STD-850 (Reference 19), allows for slouch,

and is the datum point from which the aircraft station geometry is constructed. The design eye position is a fixed point in the crew station, and remains constant for pilots of all stature via appro- priate seat adjustment.

* Horizontal Vision Line

A reference line passing through the design eye po- sition parallel to the true horizontal and normal

cruise position.

"18. Military Standard, MIL-STD-1333, AIRCREW STATION GEOMETRY

FOR MILITARY AIRCRAFT, Department of Defense, Washington,

D. C.

19. Military Standard, MIL-STD-850, AIRCREW STATION VISION

REQUIREMENTS FOR MILITARY AIRCRAFT, Department of Defense,

Washington, D. C.

33

i 1.
Design eye position

Horizontal 13 in.

vision line

S~130 desired

i minimum

back angle

Thigh tangent 31.5 in.

line / •

Back tangent

I >/l line

90o0'/-Buttock

J• reference

; line

/ / --Seat reernc

100 minimum I

/20" maximum! .5.75 in.

,f,, or heli- H

"copters, Buttock reference point planes IN 51 minimum

"/ for others Heel rest line

(Not necessarily the floor)

Figure 4. Seating geometry. (From Reference 18)

• Back Tangent Line

A straight line in the midplane of the seat passing

tangent to the curvatures of a seat occupant's back

-when leaning back and naturally compressing the back

cushion. The seat back tangent line is positioned

13 in. behind the design eye position measured along

a perpendicular to the seat back tangent line.

• Buttock Reference Line

A line in the midplane of the seat parallel to the

horizontal vision line and tangent to the lowermost

natural protrusion of a selected size of occupant sitting on the seat cushion.

34•

L.i
0 Seat Reference Point (SRP)

The intersection of the back tangent line and the

buttock reference line. The seat geometry and location are based on the SRP.

9 Buttock Reference Point

A point 5.75 in. forward of the seat reference point

on the buttock reference line. This point defines

. the approximate bottom of an ischial tuberosity, thus, representing the lowest point on the pelvic

structure and the point that will support the most

load during downward vertical loading.

o Heel Rest Line

The reference line parallel to the horizontal vision

line passing under the tangent to the lowest point

on the heel in the normal operational position, not

necessarily coincidental with the floor line.

STRUCTURAL TERMS

* Airframe Structural Crashworthiness

The ability of an airframe structure to maintain a

protective shell around occupants during a crash and

to minimize magnitudes of accelerations applied to

the occupiable portion of the aircraft during crash

impacts.

* Structural Integrity

The ability of a structure to sustain crash loads

without collapse, failure, or deformation of sufficient magnitude to: (1) cause injury to personnel,

or (2) prevent the structure from performing as intended.

o Static Strength --

The maximum static load that can be sustained by a

structure, often expressed as a load factor in terms

of G (see Load Factor, Section 2.4).

@ Strain

The ratio of change in length to the original length

of a loaded component.

35
• Collapse

Plastic deformation of structure to the point of loss

of useful loJ.-4-carrying ability. Although normally

considered detrimental, in certain cases collapse can

prove beneficial as a significant energy-absorbing

process, maintaining structural integrity.

* Failure

Loss of load-carrying capability, usually referring

to structural linkage rupture.

* Limit Load

In a structure, limit load refers to the load the

structure will carry before yielding. Similarly, in

an energy-absorbing device, it represents the load

at which the device deforms in performing its function.

"* Load Limiter, Load-Limiting Device, or Energy

Absorber

These are interchangeable names of devices used to

limit the load in a structure to a preselected value.

These devices absorb energy by providing a resistive

force applied over a deformation distance without

significant elastic rebound.

"* Specific Energy Absorbed (SEA)

The energy absorbed by an energy-obsorbing device or

structure divided by its weight. SEA is usually presented in inch-pounds per pound.

* Bottoming

The exhaustion of available-stroking distance -accompanied by an increase in force, e.g., a seat stroking in the vertical direction exhausts the available

distance and impacts the floor.

* Bulkhead

A structural partition extending upwards from the

floor and dividing the aircraft into separate compartments. Seats can be mounted to bulkheads instead of the floor if sufficient strength is provided.

36

iL i~iii
3. PRIMARY DESIGN CONSIDERATIONS

3.1 INTRODUCTION

Occupant protection and survival in aircraft accidents should

be a primary consideration in the design, development, and

testing of aircraft seats and litters. All operational requirements as specified in other design guides should also be

met. Adequate occupant protection requires that both seats

and litters be retained generally in their original positions

within the aircraft throughout any survivable accident. In

addition, the seat should provide an integral means of crash

load attenuation, and the occupant's strike envelope should be

delethalized.

3.2 OPERATIONAL ENVIRONMENT

Several environmental and operational factors other than those

associated with crashworthiness affect the design of an adequate seating system. Because of their importance in overall

design, these factors are mentioned briefly prior to the more

detailed presentation of in'ormation concerning crashworthiness.

3.2.1 Comfort

The comfort of an aircraft seat is a safety-of-flight factor

rather than a crash-safety-design factor. An uncomfortable

seat can induce pilot fatigue in a short period of time. Pilot

fatigue is an indirect cause of aircraft accidents. Comfort

is thus of primary concern and must not be unduly compromised

to achieve crash safety.

Comfort is influenced by several factors, including the vi'brational environment. Adequate comfort also involves maintenance

of adequate body angles and load distributions. Therefore,

thigh tangent angles and seat back angles are influential in

body comfort. If the back angle is less than 13 degrees, the

occupant's back will be required to counteract too much forward

moment resulting from the weight of the body acting through

centers of gravity forward of the spinal column. As the back

angle is increased beyond 13 degrees, the center of gravity is

moved back and the moment is reduced, which provides for much

greater comfort. If the thigh tangent angle is too low, too

much effort will be required to maintain the lateral orientation of the legs. If the cushion supports the lateral position

of the legs, comfort will be improved. Also, increasing the

thigh tangent angle seems to rotate the pelvis to the rear,

effectively moving the center of gravity aft and providing a

rearward moment in the pelvis that reduces the forward moment

1 37

4- -. - - -- -. - _ __ _ __ _ _ __ _ _ - .-
I

on the spine. A thigh tangent angle of 5 to 20 degrees is required by MIL-STD-1333 (Reference 17); however, it is recommended here that tangent angles greater than 10 degrees be

used to maximize comfort and to reduce submarining tendencies.

Another aspect of comfort includes the width of the seat. Too

narrow a seat can exert lateral forces on the sides of the body

or force the body to be held forward out of the constraints of

the seat bucket, again increasing discomfort. Maximum seat

widths should be provided consistent with the space availabl.

in the aircraft, including consideration for the volume around

the seat needed for lateral deflection during crash stroking

and for items such as the collective control. Typically, the

seat pan is required to be at least 18 in. wide with human factors specialists requesting 20 and compromising at 19 in. This

dimension need not be restrictive if proper consideration is

given to providing sufficient room for the seat during the initial design of the surrounding aircraft structure.

The surface upon which the occupant sits has a major influence

on comfort. The function of this surface is to spread the contact load over the largest possible area, thereby decreasing

high pressure points and preventing restriction of blood flow

in these areas. In the past, this has been accomplished by

nets or by extremely thick, soft cushions. Although such solutions provided comfort for prolonged flights, this practice is

no longer acceptable since the low spring rates of these nets

or cushions make them extremely hazardous in crash situations.

The low spring rates allow large relative velocities to build

up between the occupant and the airframe or seat during the

imposition of decelerative loads and increase the hazard to

the occupant. Thus, the cushion must provide adequate distribution of loads but not allow excessive motion during crash

loading.

Another aspect of comfort is thermal ventilation. The thermal

ventilation requirement for seat cushions is particularly important in hot, humid climates. The close contact between the

buttocks or the back and the interfacing cushions can result

in an elevation of temperatdresýcoincident with collection-of

moisture through perspiration. Provisions should be made for

air circulation to carry the hot, humid air out of this interface area, or thermal comfort will be inadequate.

3.2.2 Seat Adjustments

Passenger seats are not usually adjustable; however, in most

cases, adjustment is mandatory for crewseats. First, the cockpit and crew station have been designed for a particular eye

position. This eye position is associated with the size of a

38
* • 50th-percentile male occupant; consequently, occupants of

smaller or larger stature may not be located efficiently if

seat adjustment is not provided. Theoretically, the seat adjustment enables each occupant to adjust his eye position to

the optimum point. Typically, a ±2.5-in, vertical adjustment

from the neutral seat reference position is required tc account

for this variation in occupant size. Plus or minus 2.5 in. of

fore-and-aft adjustment is also required to permit the desired

repositioning of the eye and for locating the occupant at the

proper distance from controls, pedals, etc. Of course, human

factors should be considered in the design of adjustments.

Adjustment mechanisms should be easily found, easy to use, and

required adjustment motions should be precise, allowing the

occupant to easily get into the most comfortable position without a great deal of distraction. Further, there should be an

efficient verification that the seat is firmly locked into the

chosen position.

3.2.3 Vibration Damping

By its basic nature, the helicopter includes a great number of

vibration sources, primarily as a result of the relatively

great number of moving parts. Typical critical frequencies include the multiples associated with numbers of blades and

rotor speed. Critical conditions are located at multiples of

the main rotor speed; for example, one, two, four, and eight

per revolution. On four-bladed main rotors, the four-perrevolution frequency is typically between 18 and 20 Hz. This

driving frequency will be present constantly during cruise;

therefore, it is highly desirable that the resonant frequency

of the seat, both empty and occupied, fall outside the 18- to

20-Hz frequency range. Other frequencies, such as eight per

revolution, can also be a problem. For startup and shutdown

conditions, the resonant frequency of the seat should be high

(not lie in the range of 2 to 25 Hz), and considering the

eight-per-revolution frequency it would be desirable to keep

the natural frequency above 40 Hz. I

Seat vibrational problems are often difficult to solve because

the predetermined size and general structure of the seat seem

to control the occupied seat natural frequency rather than-the

design options that lie withi-r the limits of weight and cost.

However, the occupied seat natural frequency must be considered

since seat vibration can be very distracting to the occupant,

for example, in the lateral direction where the thighs touch

the sides of the bucket.

Stiffening of the structure is extremely costly in weight; however, in certain situations it may be the only viable solution

39

-. - - * - - - . - - - - - - - *
to the problem. Dampers that can be added to the seating system normally consist of sprung and damped masses. Usually

these mechanisms are very heavy and resorting to their use is

not acceptable in a production aircraft. Isolation of the seat

components by dash pots or elastomeric bearings may provide

possible solutions to this problem.

To summarize, consideration must be given to the vibrational

characteristics of the seat as associated with the vibrational

environment produced in the specific aircraft for which the

seat is-being designed.

3.3 CRASH ENVIRONMENT

3.3.1 DXnamics and Kinematics

When an aircraft crashes, any number of loading combinations

can be imposed on the seat. This is true for rotary- or fixedwing aircraft. It would not be useful to try to identify each

and every loading combination; however, studies that have been

completed indicate combinations of loadings that must be dealt

with in the design of the seat and restraint system. For example, the stall-spin accident typical of light fixed-wing aircraft can produce high lateral loadings, the resultant of which

can be oriented in any direction in the longitudinal-lateral

or yaw plane. Studies of helicopter crashes show very high

incidences of side impacts or rollover after impact for some

classes of helicopt,.- (Reference 20).

As an example of the dynamics and kinematics of an aircraft

crash, consider one of the new generation helicopters crashing

in a nose-up, or flare orientation. The tail boom may strike

the ground first, followed by rotation of the aircraft around

a pitch axis, then, by impact of the fuselage. The gear will

strike the ground, and, if it is a wheeled landing gear, the

tires will begin to flatten, absorbing a small amount of energy. When the rim contacts the ground, the wheel may fail as

the lower oleo strut begins stroking. After completion of the

lower oleo stroke, the second stage will begin, and energyabsorbing stroke will continue until the fuselage impacts the

ground. If the ground is relatively soft, the ground will deform under the loading of the wheels and absorb some energy.

As the fuselage impacts, the softer ground will deform again

while the fuselage structure is deforming. As the fuselage

structure deforms, additional energy is absorbed. At this

20. Haley, J. L., ANALYSIS OF EXISTING HELICOPTER STRUCTURES

TO DETERMINE DIRECT IMPACT SURVIVAL PROBLEMS, U. S. Army

Board for Aviation Accident Research, Fort Rucker, Alabama, 1971.

40

4 - __________ - -
point in the sequence, the loads can achieve the significant

magnitudes required to initiate energy-absorbing stroke of the

seat. The landing gear are designed to stroke at a lower load

than that required to activate the vertical energy-absorbing

system in the seats; thus, stroking of the gear will occur

prior to vertical stroking of the seat. This will typically

result in energy-absorbing stroke of the gear followed by an

increase in fuselage loading when the fuselage impacts the

ground and begins to crush. During some part of the crash sequence, the seat and fuselage may be stroking together. The

decelerative loads may increase and the fuselage will eventually be stopped. Depending on the conditions of the particular crash, the seat may go on stroking, until it either absorbs the residual energy of the supported mass or bottoms at

the end of its stroke. Thus, the seat may be the last item in

the load path of interest to remain in motion during the crash

sequence.

One important point here can be used to advantage by the seat

designer. In a crash with combined loading, extremely high

longitudinal or lateral loads can be applied to the seat after

stroking of the energy-absorbing gear and during fuselage

crushing. However, once the fuselage has come to a stop, crash

loading is no longer exerted on the seat, and it may continue

its stroke until either the residual energy or the seat stroke

is expended. This can be important to the designer. For example, consider a seat design that includes only vertical

energy-absorbing stroke. The seat is not required to withstand the high combined loads throughout its complete vertical

stroke, only that portion of the stroke while the lateral/

longitudinal crash loading is applied.

For those aircraft using wells, or depressions, in the floor

under the seat to provide for increased stroke distance beyond

the distance available between the seat pan and the floor, the

seat must be guided sufficiently to clear the sidewalls of the

well to utilize that additional distance. In a seat with a

low lateral spring rate, the seat may move laterally to the

point where it no longer lines up with the well under the seat

pan during the application of the longitudinal/lateral loading.

If the longitudinal/lateral loading is removed soon enough,

the seat may be able to return to alignment and still stroke

into the well under the seat. However, this occurs only if

the longitudinal/lateral loading (in certain cases) has produced elastic, rather than plastic, deformation. If the deformation has been plastic, removal of the load will not cause

the seat to return to its original over-the-well position but

will allow it to continue its vertical stroke in the deflected

configuration. On the other hand, if the elastic deformation

is not damped sufficiently, or if the distance above the well

41
is not sufficient, the rebound of the seat may carry it beyond

the well on the other side without sufficient time to return

to center as it goes through the floor plane. These motions

can be considered during seat design and development phases to

minimize a seat's weight while providing the cra-hworthy per- formance desired.

Several factors should be considered during the design of a

seat that uses a well to increase available stroke. First, as

much clearance as possible should be left between the outside

of the seat pan and the inside of the well. This will allow

for considerable deflection from the no-load position without

creating impact or interference hazards. The next consideration is that the seat be made as stiff as possible in the lateral direction to limit the extent of deflection without

imposing too high a weight penalty. Designing a seat with

energy-absorbing stroke in the lateral direction may compromise the all-important vertical stroke. Usually, the confines

around a pilot/copilot seat, which consist primarily of collective controls and consoles, do not permit sufficient lateral

motion of the seat to avoid hazardous interference with the

vertical stroke. Since the vertical stroke is the only required energy-absorbing stroke, its blockage will significantly degrade the degree of seat crashworthiness. Additionally, recent studies indicate a high frequency of thorax and head injuries (Reference 21). Allowing the seat to move either laterally or longitudinally any more than necessary could increase

the risk of head or chest impact on surrounding structure.

One could infer from the above discussion that energy-absorbing strokes in the lateral or longitudinal directions are not desirable and serve to increase the overall hazard to the occupant. However, this general statement cannot be made, as the

degree of hazard or benefit will depend on the configuration of the specific aircraft and the location of the seat within

the aircraft. In certain aircraft, space will be available

for seats that stroke in more than just the vertical direction,

and, when it is available, it may be advantageous to include

it in the system design.

Because of the cabin location of troop seats, -they typically _ have a less lethal area surroundiig, them than crewseats, and

21. Singley, G. T., III, and Desjardins, S. P., CRASHWORTHY

HELICOPTER SEATS AND OCCUPANT RESTRAINT SYSTEMS, in 0pr&

tional Helicopter Aviation Medicine, AGARD Conference roceedings No. 255, North Atlantic Treaty Organization, Advisory Group for Aerospace Research and Development,

Neuilly sur Seine, France, May 1978.

42
do not have to stroke into wells. Troop seats are typically

load limiting in the S'- longitudinal and lateral as well as vertical directions. This three-dimensional load limiting reduces

occupant decelerative loading and the crash loads on the seat

structure in the transverse direction in comparison to a

vertical-only load-limiting seat. Lower loading of the seat

allows a lighter seat design. In the case of a side-facing

seat, load limiting along the seat's lateral axis is necessary if the occupant decelerative loading during the specified aircraft forward crash impact conditions of Volume II is to be

kept within human tolerance limits for lateral decelerations.

in reviewing the dynamics and kinematics of crashing aircraft,

it beoomes quite apparent that all combinations of orientations,

loading, and load directions can exist. (Volume II presents

a detailed discussion of crash impact dynamics and kinematics.)

It should also be remembered that the seat is designed to absorb only a portion of the crash energy required to decelerate

the occupant in a tolerable environment. There are numerous

crash orientations in which the aircraft has a lateral component of impact velocity, whether it results from a lateral

drift of the aircraft or from its attitude at impact. These

components of velocity can produce high lateral loading of gear, which, in some cases, may simply break off before absorbing

significant energy. Consider, for example, the case of an aircraft impacting the ground with a high roll angle. Loss of

gear will result in the aircraft fuselage impacting the ground without the reduction in energy normally attributed to stroking

of the gear. Therefore, systems analyses must take this factor

into account. As an example of the possible dangers, it might be decided that landing gear should absorb all the crash energy

associated with the 42-ft/sec vertical impact; therefore, seat

stroking would not be required. The results of applying this

logic to hardware would seriously reduce the overall crashworthiness of the aircraft in those crashes where the full energy absorption of the gear could not be realized. Therefore,

seats should contain the minimum energy-absorbing stroke defined in this document, regardless of the energy-absorption

capacity of the gear.

After- a helicopter crashes, the rotating main rotor may strike

the -ground or other obstacles and roll the helicopter onto its

side. Because of the high center of gravity, the helicopter

may roll over without any added lateral impulse from the main

rotor blades after gear failure. In any case, the kinematics

of crashed helicopters can be quite complex and violent, and

the helicopter may come to rest in any orientation. Because

of these kinematics, loads are specified in all directions for

seats. This subject will be covered in more depth later in

this volumey however, the crash kinematics of these aircraft

I 43

t TF•-'i' i• r-•-at I I iil.1
demand strength requirements in all directions, including upward and aftward. In this regard, it should be remembered that

the seat may have used a significant portion of its available

vertical stroking distance durinc, the major impact. If the

aircraft should then follow through with a flip, or land on

its back, the system should maintain the seat near its final

stroked position rather than allowing the seat to return to

its original position. Upward travel could be extremely hazardous if the roof of the fuselage were severely crushed and

the occupant were free to travel unrestrained back toward his

initial position. Severe head and/or neck injuries could resuit.

In summary, it must be remembered that, to produce a crashworthy design, systems analyses must consider likely combinations of loadings, including potential losses of energyabsorbing structure, such as landing gear.

3.3.2 Design Conditions and EnveloP2s

The design impact conditions for light fixed- and rotary-wing

aircraft are presented in Volume II and are repeated here in

Table 2. All seats, restraint systems, and litters should be

designed for these impact velocities and provide the desired

performance in the design crash environments.

TABLE 2. SUMMARY OF DESIGN CONDITIONS

FOR ROTARY- AND LIGHT FIXEDWING AIRCRAFT

Velocity

Impact change Direction (ft/sec)

Longitudinal 50

Vertical 42

Lateral* 25

Lateral** 30

*Light fixed-wing, attack, and cargo

helicopters.

**Other helicopters.

44
3.3.3 Structural Distortion

4- Structural distortion of the airframe and its resulting loading of the seat must be considered in the design stages. For

example, a ceiling-mounted seat may experience lower loads than

a floor-mounted seat because of the distortion or deflection

of the roof and supporting walls. However, additional stroke

distance may be required due to the inefficiency of the stroke

provided by distortion of the airframe as compared to that provided by a load-limited seat. The effective, stroke of a seat

considered to be rigidly attached (no energy absorbers between

the seat and roof) to the roof also must be ýconsidered. If

the seat pan is 12 in. from the floor of the aircraft and the

roof of the aircraft is expected to distort downward on the

order of 12 in., careful consideration must be given to eliminating rebound rather than increasing total stroke, which could

result in bottoming. In the practical case, the roof probably

distorts something less than the distance between the seat pan

and the floor of the aircraft; therefore, energy-absorbing

stroke should be provided in the seat to maximize usage of the

available space. A systems analysis should be applied to this

situation to establish the correct combination of variables.

A considerable amount of the downward motion of an aircraft

ceiling may be elastic. It would be advantageous to eliminate

the rebound from this elastic distortion from the occupant and

seat. Consideration could be given to a device that allowed

vertical downward motion of the seat but restrained it from

following the roof during its elastic rebound. It is possible

to think in terms of an energy-absorbing device under the seat,

connected to the floor, that resists motion in both the downward and upward directions. This would allow only partial return of the seat resisted by the energy-absorbing system.

Another alternative would be to provide energy absorbers between the seat and the roof that would stroke when the roof

returned to its equilibrium height. Again, a device that locks

the seat in its lowest position would be required with this

concept. Adequate support of the ceiling to support the applied loads with low deflections eliminates the prpblems mentioned. Efficient use of ceiling-mounted seats can be achieved

in these aircraft.

A major consideration in providing crashworthy seating systems

is the possibility of a local distortion in the part of the

aircraft to which the seat is attached. For example, a floormounted seat may have to withstand severe distortions as a result of underfloor and floor deformations caused by impact

forces. If the aircraft crashes on uneven ground or encounters

rocks or stumps, distortions of the underfloor structure can

occur. The seat structure or seat attachment to the floor

2* 45

&
should be adequate to permit these distortions without producing failure of the seat structure or its attaching mechanisms.

It should be noted that the forces causing this distortion cannot be resisted by the seat structure. In other words, it is

not feasible to build a seat strong enough, if rigid, to maintain the attachment to the aircraft in these situations. The

crash loads causing the distortion will, in most cases, exceed

any strength that can be designed into the seat, thus, produc- ing failures if not adequately accounted for in the design.

Likewise, distortion of bulkheads in bulkhead-mounted seats

presents the same type of problem. It is apparent that local

distortion of a bulkhead usually will not be of the magnitude

of the distortions that can occur in the floor structure of an

aircraft. Rocks and stumps can produce extremely large local

deformations of structures upon which floor-mounted seats are

mounted. However, rocks and stumps normally will not be involved in distortion of bulkheads and bulkhead structure. Consequently, the distortion requirements for seat mountings on

bulkheads are less severe. A recent search of USASC crash records identified no known cases of bulkhead-mounted seat loss

due to bulkhead distortion or fracture of attaching structure.

It is expected that sidewalls will deform more than transverse

bulkheads, although they would not be as susceptible to rocks

and stumps as floors. The deformation would usually be one of

the wall buckling outward near the floor and changing the lateral and vertical relationships between attachment points.

However, it should be remembered that in helicopters, sidewallmounted seats are not usually pilot or copilot seats and therefore, are usually not of the stiffness that would create a

problem in the environment described. For some fixed-wing aircraft, though, this may not be the case. For these cases, the

aircraft/seat interface should be designed to be compatible

by allowing flexibility in the seat, in the attachments, in

stiffening the sidewall of the aircraft, or by simply not attaching rigid seats to sidewalls. Floor, bulkhead, and sideward warpage requirements are presented in Section 4.4.5, Joint

Deformation.

3.4 APPLICABILITY OF CRITERIA

The recommendations in this volume will apply to all categories

of U. S. Army aircraft. Those recommendations having application to a specific class or category of aircraft only will be

so indicated by the text.

46
I

S3.5 ACCEPTANCE CRITERIA

In addition to operational requirements specified in other design guide documents, seats and litter systems should be designed to provide occupant protection under crash conditions

as specified in Volume II. Appropriate stress analyses, tests,

and operational requirements outlined in this volume should be

met by every seat, restraint, litter system, and by the cockpit and cabin interior prior to acceptance.

3.6 SELECTION CRITERIA

Crashworthy seats, restraint systems, litter systems, and cockpit and cabin materials should be evaluated on the basis of

the occupant protection provided, and on their anticipated reliability and serviceability under the oierational and potential crash conditions expected.

47.

II

~.I

O 47

_.•°•.
1 4 .' ~ ~ ~ . *~l.. . "

4. DESIGN PRINCIPLES FOR SEATS AND LITTERS

4.1 INTRODUCTION

There are several types of Army aircraft seating systems: pilot, copilot, crew chief, gunner, observer, student, medical

attendant, troop, and passenger. Cockpit seats are typically

forward-facing; however, cabin seats may face in any direction.

Most are single-place seats, but in a few aircraft, two-,

three-, and four-occupant cabin seats are provided. A single

occupant seat is the preferred configuration in order to avoid

situations where the energy-absorbing systems of multi-unit

seats are rendered ineffective due to less than full occupancy

(insufficient weight to activate the energy-absorbing mechanisms at loads within human tolerance limits). To the maximum

extent practical, seats should be interchangeable to enable

standardization. It is desirable that all seats face in the

same direction so that the seat backs protect occupants from

loose equipment which can become flying projectiles during.

crash impact.

The rearward-facing seat is optimal for providing maximum support and contact area in longitudinal impacts. The only critical impact sequence for the rearward-facing seat is one that

involves a severe lateral component that allows sideward movement of the occupant prior to application of the longitudinal

or vertical pulse. However, lateral torso movement can be prevented by use of an adequate restraint system of much lighter

weight than that required for other seat orientations. When

practical, the rearward-facing seat should be used.

Those crew members required to face forward in the conduct of

their duties can be afforded adequate protection by the use of

a restraint system consisting of shoulder straps, a lap belt,

and a lap belt tiedown strap as discussed in Chapter 7. The

lap-belt-only restraint is undesirable, as noted in the human

tolerance section of Volume II. If all forward-facing passengers are provided with adequate upper- and lower-torso restraint, forward-facing seats -re acceptable as a second choice

to rearward-facing seats. If a single, diag6nal, upper-torso

restraint is used, it should be placed over the outboard shoulder of the occupants to provide restraint against lateral pro- trusion of the occupant outside the aircraft or impact with the

sidewall.

Previously, side-facing seats have been provided with lap belt

restraint only. This arrangement is considered completely inadequate for providing crash protection. Even with the addition of a shoulder harness or diagonal chest strap, the tolerance to abrupt acceleration is minimal. The use of side-facing

48
seats is least desirable from the crash safety standpoint;

however, when no reasonable alternative to their use exists,

adequate restraint must be provided. If a single, diagonal,

upper-torso restraint is used, it should be placed over the

forward-facing shoulder (relative to the aircraft).

4.2 LITTERS AND THEIR ORIENTATION

The supine position of a litter patient is ideal for resisting

vertical impacts. The contact arep is the maximum possible,

and the decelerative forces act transversely to the body. tor

existing litters, the major problem occurc as a result of impact forces in the lateral/longitudinal plane. The relatively

flat litter surface makes it difficult to provide an adequate

restraint harness to resist these loads. The current practice

of wrapping two lengths of webbing around the litter offers a

degree of restraint oriented transversely to the body, but only

frictional forces prevent the body from sliding off the litter

in the lengthwise direction.

Litters should be installed laterally to provide more positive

restraint for expected combined crash forces.. A lateral litter

orientation also will prevent the litter from becoming completely detached from its current supports as occurs in a longitudinal orientation explained in Reference 22. The litter

must withstand all of the environments previously described

for the seats.

4.3 MATERIALS

Designers should select materials that offer the best strengthto-weight ratios while still maintaining sufficient ductility

to prevent brittle failures. The guidelines in this section

will alert the designer to certain material properties that

can contribute to improved structural designs. These properties include ultimate strength, elongation, and energyabsorbing capabilities. The standard method for selecting materials using elastic analysis is adequate for most conditions

in the working lifeof an article. For crashworthiness, however, only one applicat-i-on of the maximum load is expected,

and the behavior of the material beyond the yield point generally is important.

22. Weinberg, L. W. T., AIRCRAFT LITTER RETENTION SYSTEM DESIGN CRITERIA, Aviation Crash Injury Research (AvCIR), Division of Flight Safety Foundation, Inc.; USAAVLABS Technical Report 66-27, U. S. Army Aviation Materiel Laboratories, Fort Eustis, Virginia, April 1966, AD 632457.

49

S_...t.. . . .
The degree of ductility needed in a seat's basic structural

parts is highly dependent upon whether the seat structure is

designed to absorb energy by the use of a separate Ilcdlimiting device or whether large plastic deflections of the

basic structure are required. As a general rule, a value of

10-percent elongation is a rough dividing line between ductile

and nonductile materials. The 10-percent value is recommended

as a minimum for use on all critical structural members of

nonload-limited seats because the exact peak load is unpredict- able due to pulse shape, dynamic response of the system, and

velocity change. A minimum elongation of 5 percent in the

principal loading direction is suggested for use on critical

members of load-limited seats because the loads and strains

are more predictable.

Castings are not recommended for use in primary load paths.

In general, their quality is more difficult to verify and reproduce, and their ductility and fracture toughness are less

than for forgings.

The effects of stress corrosion, (for example, selection of

7075 aluminum alloy in a T73 condition rather than T6), must

be considered, as well as hydrogen embrittlement due to heat

treating or various processing steps such as pickling (for example, 17-4PH stainless steel). In short, adherence to all

the normal engineering design principles must prevail.

Flammability and toxicity retardation requirements are discussed in Volume V. Upholstery padding and other materials used

in seats should meet the specified requirements.

4.4 STRUCTURAL CONNECTIONS

4.4.1 Bolted Connections

For the manufacture of basic aircraft structure, most aircraft

companies recommend 15- and 25-percent margins of safety for

shear and tensile bolts, respectively. These factors are intended to allow for misalignment of holes, stress concentrations, and fatigue. Fatigue is not generally a factor in the

design of a seat or litter system fitting, since high loading

of the fitting would be a one-time situation. Therefore, the

safety factor for shear and tensile bolts located in loadlimited portions of the seat where loads can be predicted accurately, can be reduced to 5 and 10 percent, respectively.

Also, good aircraft engineering practice dictates that bolts

less than 0.25 in. in diameter should not be used in tensile

applications because of the ease with which these smaller bolts

can be overtorqued. Because of the obvious advantages of

structure being able to distort while maintaining load-carrying

50
ability, fasteners of maximum ductility for the application

should always be selected. Where possible, fasteners such as

bolts and pins should have a. minimum elongation of 10 percent in the longitudinal and transverse directions.

4.4.2 Riveted Connections

The guidelines for riveted joints are presented in MIL-HDBK-5,

and it is recommended that these guidelines be followed (Reference 23).

4.4.3 Welded Connections

Welded joints can be 100 percent efficient; however, the actual efficiency is dependent upon the skill of the welder, the

process used, and the inspection procedures followed. Welded

joints can be completely acceptable and even superior to bolted

or riveted joints. However, strict inspection procedures

should be used to ensure that welded joints are of good quality. Welded joints may result in stress concentrations and

misaligned parts in a manner similar to bolted joints; therefore, the cross-sectional area of the basic material in the

vicinity of a welded joint should be 10 percent greater than

the area needed to sustain the design load. Welding processes

are discussed in Military Specifications MIL-W-8604, -6873, -45205, anrl -8611; these specifications should be used as

guides to ensure quality welding.

4.4.4 Seat Attachment

Cockpit seats are either bulkhead or floor mounted. Acceptable

means of attaching seats to the cabin interior are listed below

(refer to Section 3.3.3 for a discussion of ceiling-mounted

seats and ceiling support stiffness):

1. Suspended from the ceiling with energy absorbers,

and wall stabilized.

2. Suspended from the ceiling with energy absorbers, and floor stabilized.

3. Wall mounted with energy absorbers.

23. Military Handbook, MIL-HDBK-5, METALLIC MATERIALS AND

ELEMENTS FOR ASROSPACE VEHICLE STRUCTURES, Department of

Defense, Washington, D. C.

51
4. Floor mounted with energy absorbers.

5. Ceiling and floor mounted (vertical energy absorbers

above and below seat).

Suspension or mounting of all seats should not interfere with

rapid ingress or egress. Braces, legs, cables, straps, and

other structures should be designed to prevent snagging or

tripping. Loops should not be formed when the restraint system

is in the unbuckled position. Cabin seats must often be designed so that they may be quickly removed or folded and secured. Tools should not be required for this operation. The

time required by one person to disconnect each single occupant

seat should not exceed 20 sec. The time required by one person

to disconnect multi-occupant seats should not exceed 20 sec *

multiplied by the number of occupants. All foldable seats

should be capable of being folded, stowed, and secured or unstowed quickly and easily by one person in a period not to exceed 20 sec multiplied by the number of occupants.

4.4.5 Joint Deformation

Floor distortions as a result of impact can cause failure of

the seat structure or tiedown connections in an aircraft crash

(see Figure 5). A floor distortion can take the form of a

bulge or dish in the floor surface between the seat tiedown

connections. This produces a rotation of the seat relative to

the floor surface, resulting in a connection failure if the

deflection limits for the attachments are exceeded. A twisting or warping of the floor surface can also take place, producing distortion loads in the seat structure. Seat or connection failure can result from the additional loads imposed. I

The seat designer must anticipate possible floor bulging or

warping and take appropriate measures in seat structural design to minimize the adverse effects.

For basically rigid seat structures that are distorted the critical design parameter appears to be the torsional rigidity of

the seat pan, bucket, and/or structural members. If the torsional rigidity is low, only small forces-are introduced. However, for stiff seat members, the warpage forces may produce a

structural failure or impose a preload that, when coupled with

-crash inertial loads, results in failure. A high torsional

rigidity in the seat pan may arise from integrating stiff lateral cross tubes between side trusses so that the tubes must

also twist with the seat pan. Consequently, it may be desirable to connect the cross tubes to the seat pan in such a way

that the seat pan is free to twist independently of the cross

tubes or to design the crossmembers to be soft in torsion. Integrally armored crewseats are stiff and difficult to release

52
r.r

* i

Distortion in Final position

lateral direction

(around roll

Seat mount locations

, •%,. Distortion in.

S "'" ' fore/aft (around pitch direction axis)

Rock, eart

or other

obstacle Initial position

Figure 5. Sketch illustrating buckling

or "dishing" formation.

from the support structure in order to permit distortion. One

method used successfully to solve this problem has been a threelegged seat. The three support points can follow the floor

movement without distorting the seat structure because the seat

is free to tip (Reference 24).

24. Desjardins, S. P., and Harrison, H., THE DESIGN, FABRICATION, AND TESTING OF AV INTEGRALLY ARMORED CRASHWORTHY

CREWSEAT, Dynamic Science, Division of Marshall Industries; USAAMRDL Technical Report 71-54, Eustis Directorate, U. S. Army Air Mobility Research and Development

Laboratory, Fort Eustis, Virginia, January 1972, AD742733.

1 53
To prevent seat connection failures induced by floor dist ion,

structural joints should be capable of large angular displacements in all directions without failure. A seat designed prop- erly for structurally integral load limiting would also satisfactorily accommodate floor buckling and warping under crash

conditions. Figure 6 illustrates recommended limits of floor

warping or buckling that must be withstood by all floor-mounted

seat designs. The mounts should be capable of withstanding a

')-degree warp of the floor, as well as a ±10-degree rotat.on

"o. ut a roll axis of a single track. The angles are based on

Jistortions that have been noted in potentially survivable accidents.

With respect to the floor surface and to accommodate rotations

that result from floor bulging, several design configurations

may be considered. Two of these are presented below and are

illustrated in Figure 7.

• A deliberate plastic hinge of sufficiently ductile

material may be incorporated into the tiedown connection design. This plastic hinge would be required

to permit yielding without failure up to a rotation

angle that exceeds the maximum anticipated as a result of floor bulging. The hinge also would be required to carry the associated compressive, tensile,

and shear loads in order to retain the seat while

yielding in bending.

* A structural release such as a ball-and-socket joint

may be used to pcmit relative rotation.

Other methods, such as a combination of a plastic hinge about

one axis and rotation about an axle or pin oriented along a

perpendicular axis, are acceptable also. The joint must be

capable of sustaining large tension, compression, and shear

forces during and after rotation.

The effect of not providing for relative seat leg-to-floor rotation can be illustrated by an actual example. The rear-.legs_- of a crewseat on early models of a7 U. S. Army helicopter were

attached to a base frame with castings-as illustrated in Figure 8. These castings failed repeatedly in accidents as a result of combined axial and bending stresses acting at the region of stress concentration. Studies showed that the seat

could sustain a longitudinal decelerative force nearly twice

as great when the bending moment at the juncture between the

rear leg and the track fitting was removed.

54

4 . - - - _____ , __ - - . . .__________________- ...-- • •-
I

S- Thi +-i0* 4oiiaini mi5 lutae nFgr . Temmn a

°z

....... in

_Y

x

Figure 6. Static test floor warpage requirement to improve

the probability of seat retention in crashes.

This modification is illustrated in Figure 9. The moment was

relieved by cutting the corners off the casting so that only

the section around the center bolt remained. The joint was

55.

nw :
S

I'~. 

, j

Hinge, 0 i

I I

(a) Plastic hinge (b) Ball-and-socket

Joint

Figure 7. Concepts for release of floor-distortioninduced moments.

thereby 

SubsequenIt changed from a fixed- to a pinned-end configuration. tests showed improved load-carrying capacity.

Other 

spherical 

methods of relieving torsion and moments include using

For example, bearings and slotting holes through which -bolts pass. if a crossmember during floor warping, is required to move torsionally

vided for slots that relieve the leads can be pro- fasteners at end fittings. This is illustrated Figure 10. Figure in

released 11 illustrates an example of a fully- joint acted on by two torsional loads and a moment.

The same general principles also apply for that apply for floor-mounted seats bulkhead-mounted tion and degree seats except that the deflecthan that of 

of warping of the bulkhead appear to be less

being less vulnerable 

the floor. This is probably due to the bulkhead to local planar distortion caused by

I 56
Occupant O load

Aft seat leg Z • (steel tube)

A4

Castng

Floor tracks

Ar-Floor

Figure 8. Aft seat leg casting attachment.

IN Pi

Failure point

(stress con- M

centration)

Fixed joint 'Pinned joint

Floor Floor

track track

View ABefore modification view A After modification

Figure 9. Aft seat leg casting attachment modification.

-57

ti..-
Ti

Nut

TorsionT

release-\

Torsion release

Figure 10. Torsional release of joints.

items such as rocks and stumps impacted by the underfloor

structure. -A possible bulkhead distortion configuration is

shown in Figure 12. The recommended angular deflection requirement for bulkhead-mounted seats is a 5-degree rotation in

the plane of the bulkhead. To accommodate local deformation,

each attachment of the seat to the bulkhead should be released

to permit t10-degree rotations in any direction. One technique

for accomplishing this is with spherical bearings, as illustrated in Figure 13.

Sidewall-mounted seats require the same considerations as

bulkhead-mounted seats. As mentioned previously, the sidewalls

58

_ i
6TorTorsion

release

release

release

Figure 11. Fully released joint.

of aircraft tend to bow outboard during impacts-with high vertical loading. Therefore, it is advisable that these seats be

designed to accept relatively large distortions without failure. Although the angles are not known, it is expected that

they may reach 25 degrees.

Seats that are mounted totally on the sidewall should not

create a problem, as they will simply move with the sidewall.

Extremely flexible seats also should be inherently immune from

these problems. However, rigid seats mounted to both the floor

59

• .........-......... '--',-...a• .. ". •
Bulkhead

Seat attach

-[points (4 places)

Lower fuselage

(a) Initial configuration

Bulkhead

(b) Postcrash configuration

Figure 12. Bulkhead in-plane warping.

60

•lilI I 11 II'TII'IT'I
Clevis

I I

ug LugClevis

Figure 13. Universal release of a joint.

and the sidewall will require special design considerations.

One way to provide the flexibility needed is to include releases such as pin joints, oriented to allow rotation around

an aircraft roll axis. An example is shown in Figure 14. The

attachments should be designed to permit the angle 0 to reach

25 degrees at the maximum dynamic deflection.

The underfloor, bulkhead, or sidewall structure must be designed to be compatible with the seat. For example, the design

of structural releases between the seat and the track may enable the seat to maintain its attachment during large floor deformations but may add to the torsional responsibilities of underfloor beams. If a large downward load is applied to the

floor structure through a joint that does not carry moment (released), then the underfloor beams must resist any moment that

may be developed without assistance from the seat structure.

To illustrate, take the case of a seat strut attached through

a release to the front-floor tracK. During longitudinal loading in the forward direction, the strut is loaded in compression and applies a large downward load at the release. Any

eccentricity between the load vector and the centroid of the

underfloor beam will produce torsional loading around the

beam's long axis. The beam must possess the capability to resist this torsional load through either its own torsional

strength or that of its supporting structure.

61
Initial fuselage

configuration

Deflected

fuselage

configuration

• A

\ ~/,

Floor l

Figure 14. Pin joint releases oriented to allow rotation

around an aircraft roll axis.

4.5 STRENGTH

4.5.1 General

An elastic stress analysis, as used in the design of airframes

and aircraft components subjected to normal flight loads, is

inadequate for the study of all the structure in a crash situation. For normal flight loads, keeping the stresses well below

the material yield stress to avoid permanent deformation is

necessary because of fatigue problems and, perhaps, other considerations. In a crash situation, however, where only one

62
application of maximum load is expected, fatigue is not a factor, and the final configuration of a structural component or

its subsequent operational use need not be considered. Consequently, the load-carrying capacity of components deformed beyond the elastic limit should be considered in determining the

ultimate seat strength. As a matter of fact, it is advisable

for certain items in the load path to use the rupture strength

as listed for many materials in MIL-HDBK-5 (Reference 23).

The concepts of limit analysis or, in some circumstances, large

deformation analysis, may be employed to make the best use of

* materials in certain components.

It may appear that the only difference between an elastic

stress analysis and an ultimate strength analysis is that the

former is more conservative. However, a more significant distinction is demonstrated by a comparison of two designs having

the same maximum stresses for elastic behavior but decidedly different load-carrying capacities when the loads exceed the

elastic limits. For example, consider the following two similar designs: (a) two simple beams spanning three supports and

(b) a continuous beam spanning the same three supports, as illustrated in Figure 15.

Mw12

(a) Two simple beams (b) Continuous beam

Figure 15. Comparison of analysis methods

for simple beams.

For a uniformly distributed load, w, the bending moment diagrams are as shown (assuming elastic behavior). It is noted

that in each case the maximum bending moment is wl /8 and each

design has the same stress. There is a temptation to equate

the designs from a strength viewpoint. However, considering

design (a), if the load is gradually increased, the bending

moment at the center of each span will eventually equal the

moment resistance capability of the beam. For a ductile material, a yield hinge would fcrm then at these maximum moment

* 63
points. Additional load could not be accepted without a mechanical collapse. This critical load would represent a realistic ultimate capacity for the beams. On the other hand, when

a yield hinge occurs in design (b) under similar circumstances,

it would occur at the middle support and, hence, not produce a

collapsing mechanism. The load, w, could be further increased

without collapse until a second set of yield hinges forms between the supports. Only then would collapse occur. It is

intuitively evident, and may be demonstrated by analysis, that

design (b) sustains a much greater ultimate load than does design (a), yet the difference is not discernible from elastic

analysis. The design of an entire occupant retention system,

ignoring inelastic postyield behavior, would result in compo- nents of varying ultimate strengths, some much stronger than

others. The overdesigned components do not increase the

strength of the system. It is desirable that all components

work at the same allowable strength level just before failure.

A 1963 study of the restraint system used in three U. S. Army

aircraft indicated that the strengthening of a few weak links

in the tiedown chain improved the crash strength of these systems by a factor of 2 with only minor weight increases (References 25 through 27). A simple example of the benefit of

strength analysis beyond the elastic limit is the improvement

in the tiedown strength of the crewseat floor track in one of

the three aircraft. In the existing arrangement, the seat leg

may be positioned directly above a pair of seat track tiedown

bolts (Figure 16). The elongation of the bolts prior to their

failure would not be sufficient to permit bending in the floor

track; thus, no appreciable load could be transmitted to the

adjacent pair of bolts. To improve the ultimate strength of

this connection, it was suggested that aluminum collars, which

compress at a load slightly less than the breaking strength of

25. Haley, J. L., Jr., and Avery, J. P., Ph.D., PERSONAL RESTRAINT SYSTEMS STUDY - HC-lB VERTOL CHINOOK, AvCIR 62-26,

Aviation Crash Injury Research (AvCIR), Division of Flight

Safety Foundation, Inc., Phoenix, Arizona, November 1962.

26. Haley, J. L., Jr., and Avery J. P., Ph.D., PERSONAL R-ESTRAINT SYSTEMS STUDY - HU-IA AND HU-lB BELL IROQUOIS,

AvCIR 62-27, Aviation Crash Injury Research (AvCIR), Division of Flight Safety Foundation, Inc., Phoenix, Arizona, December 1962.

27. Haley, J. L., Jr., and Avery, J. P., Ph.D., PERSONAL RESTRAINT SYSTEM STUDY - CV-2 DE HAVILLAND CARIBOU, AvCIR

62-16, Aviation Crash Injury Research (AVCIR), Division of

Flight Safety Foundation, Inc., Phoenix, Arizona, April

1964.
Present attachment

Seat reaction Leg stud

0.50 in.J Floor track |

i

Proposed attachment

I

• ;•':7• •0.25 in.

Aluminum collars added I

Figure 16. Seat leg anchorage to floor track.

the bolt, be added beneath the nut. Thus, the collars would

yield prior to failure of the center bolts and permit the track

to bend and transmit some load to the adjacent bolts. This

arrangement approximately doubled the ultimate tiedown strength

of the floor track while adding a negligible amount of weight.

4.5.2 Limit Analysis Concepts

Where ductile materials are used, strain concentrations do not

produce rupture prior to significant plastic deformation. If

the geometric configuration of the structure permits only small

elastic deflections, a rigid-plastic mathematical model may be used. This permits the-use of a limit analysis, which assumes no deformation of structUre until sufficient plastic

hinges, plastic extensors, etc., exist to permit a geometrically admissible collapse mode.

Limit analysis is concerned with finding the critical load sufficient to cause plastic collapse with the physical requirements of static equilibrium, yield conditions for the materials,

65
and consistent geometry considerations. The principles of

limit analysis are well developed by a number of authors (References 28 and 29, for example). Two useful principles are

mentioned here: the upper and lower bound theorems. The upper

bound theorem for the limit load (collapse load for a rigidplastic structure) states that the load associated with the

energy dissipated in plastic deformation will form an upper

bound for the limit load. The lower bound theorem, on the

other hand, states that the load associated with a statically

admissible stress distribution, which at no poiat exceeds the

yield conditions, forms a lower bound for the limit load. Use

of the upper and lower bound theorems to bracket the limit load

for a given structure makes it possible to obtain a realistic

evaluation of the structure's load-carrying capacity.

4.5.3 Lar.e Deformation Analysis

If a structure contains elements that will permit large, stable elastic deformations when under load, the equilibrium of

the deformed state must be considered in evaluating ultimate

* strength. For example, if a suitable attachment is made to a

thin flat sheet rigidly fixed at the edges so as to load the

sheet normal to the surface, a diaphragming action will occur.

The equilibrium and stress-strain (e)astic-plastic) relations

for the deformed state would determine the load-carrying capacity. An example of this situation is a seat pan in which membrane rather than flexural stresses are important.

4.5.4 Strain Concentrations

Handbook stress concentration factors provide sufficiently accurate data to allow the designer to modify the strizcture in

the vicinity of stress concentrations. When large deformations at high load-carrying capacity are desired, as in energyabsorbing seats, these areas frequently become strain concentration points and rupture occurs, due to excessive strain, in

areas with little deformation and energy input. Large amounts

of energy can be absorbed in the structure only if large volumes

28. Beedle, L., PLASTIC DESIGN OF STEEL FRAMES, John Wiley

and Sons, New York, 1958.

29. Hodge, P. G., Jr., PLASTIC ANALYSIS OF STRUCTURES, McGrawHill Book Company, New York, 1959.

66

-A ____________________________-_
of material are strained uniformly. For further information on

the subject, see pages 69-73 of Reference 30.

4.6 RESTRAINT SYSTEM ANCHORAGE

- The design requirements for occupant restraint systems are presented in Chapter 7; however, the seat designer must consider

the effect of the anchorage of the restraint system on the characteristics of the seat design. The restraint system should be

anchored to the seat rdther than to basic aircraft structure.

If the restraint system is anchored to basic aircraft structure, a desirable reduction of loads on the seat frame results;

however, the restraint system must be designed to permit the

energy-absorbing deformation of the seat during an impact.

For example, if a load-limited seat strokes vertically and the

seat belt is anchored to the floor, the loosening of the belt

would permit the occupant to "submarine" under the belt or to

move laterally. When the harness is anchored to the seat structure, the problem of maintaining a snug harness is reduced.

An advantage of attaching the shoulder harness to basic aircraft structure is the large reduction in overturning moment

on the seat. To make this attachment acceptable, a simple

load-limiting device might be incorporated into the shoulder

harness anchorage to allow for longitudinal or vertical movement of the seat. On some aircraft, where room allows it,

another option is to locate the anchor point far enough to the

rear of the seat to allow vertical energy-absorbing stroke of

the seat with only a rotation of the strap about the anchor

point. If the distance is sufficiently large, the fore-and-aft

motion resulting from the strap swinging in an arc can also be

insignificant.

4.7 CRASH ENERGY ABSORPTION

4.7.1 General

The magnitude of a crash force is a function of the input velocity and the stopping distance. The stopping distance is

controlled basically by the crushing of the airframe and landing gear in a given direction coupled with the gouging of the

impact surface. The magnitude of the average deceleration of

30. Turnbow, J. W., et al., AIRCRAFT PASSENGER-SEAT-SYSTEM

RESPONSE TO IMPULSIVE LOADS, Aviation Safety Engineering

and Research (AvSER), Division of Flight Safety Foundation, Inc.; USAAVLABS Technical Report 67-17, U. S. Army

Aviation Materiel Laboratories, Fort Eustis, Vir ;iria,

August 1967, AD 661088.

* 67

L~..
a given point of the aircraft may be calculated from the following equation:

a 2 2 2 2 i.v -vf v -vf

i;2S or = 2gs

where a = average deceleration, ft/sec2

G a average deceleration, G

vO a initial velocity, ft/sec

vf = final velocity, ft/sec

g = acceleration due to gravity, 32.2 ft/sec2

S = total displacement of the point of the aircraft

with respect to the ground, ft

It can be seen from the equation that t ' magnitude of the deceleration is inversely proportional to the stopping distance.

In the ca!- of a rigid structure impacting a nonyielding sur-

"face, the deceleration would be infinite. Some crushing of

structure and soil educes or attenuates the deceleration to

finite levels. Often, however, there is insufficient crushing to attenuate deceleration magnitudes to human tolerance

levels. Tolerable levels can be achieved by increasing the

stopping distance. The extra stopping distance may be provided by using: (I) additional crushable airframe structure,

(2) energy-absorbing landing gear, (3) a seat design that

possesses an energy-absorption mechanism(s) (load-limiting or

controlled seat collapse), or (4) a combination of methods (1),

(2), and (3).

The energy-absorption capability of a seat structure is also of

considerable importance in evaluating the seat dynamic strength.

Due to extension of the restraint harness, compressibility of

the soft human tissue under the harness, penetration into the

seat cushion, and relative movement of body parts, the occupant's center of gravity acquires a velocity relative to the

airframe during an abrupt deceleration.

Depending upon the magnitude and duration uf the deceleration

pulse, as well as the nature of the connection between the occupant and the seat structure, '-he maximum relative velocity

may be large. The seat structure, in order to perform its intended retention function, must then either M) possess the

capability of sustaining the maximum inertial force imposed by

68
the deceleration of the occupant and the seat without collapse, "or (2) possess sufficient energy-absorption capacity to reduce

the occupant's relative velocity to zero before structural failure occurs. The first alternative may result in an excessive

strength requirement because the input pulse shape and elasticity of the restraint system and cushion can result in significant dynamic overshoot. Computer simulation and experimental

observation have shown that overshoot factors range from 1.2

to 2.0, necessitating a seat design strength requirement of

24 G to 40 G to accommodate an input floor pulse of 20 G.

The second alternative of using collapse behavior (load limiting) appears to offer the more practical approach to seat design. With this option, the seat structure would begin plastic

deformation when the acceleration of the occupant and seat mass

reaches a level corresponding to the critical structural load.

The seat must absorb enough energy without failure to stop the

motion of the occupant relative to the aircraft. Of course,

this energy must be absorbed at force levels within human tolerance limits to provide the intended protective function.

In an attempt to eliminate common misconceptions regarding the

role of energy-absorbing seats, a few introductory comments

are made:

* The seat energy-absorbing system does not absorb all

the -gy associated with the impact ce--nocty. The

ser g aeriences the total velocity change; however, mo .the energy is absorbed by deforming earth,

stroKing landing gear, and deforming structure.

o The absorption of energy by the above processes produces the triangular-shaped deceleration versus time

pulse used as the design input to the seat.

* The seat energy-absorbing stroke simply lengthens

the stopping distance of the occupant by allowing

energy-absorbing stroke of the seat to occur as the

other energy-absorbing processes are nearing completion. In a crash in which the aircraft comes to

rest in the major impactr-much of the seat stroke |

can occur after complete deceleration of the aircraft

fuselage. Thus, after the fuselage stops, the seat

may continue to stroke until its kinetic energy has

been exhausted.

* Disregarding dynamic response differences, the same

stroke distance is required to decelerate any mass at

a given deceleration magnitude. Therefore, lighter people do not require shorter strokes than heavier

69
people for the same deceleration magnitudes. Of

course, loads required to decelerate occupants of

different weights at equal deceleration magnitudes

must vary with occupant weight.

* Consideration of the first comment explains why it

is detrimental to allow slack to develop in the restraint system or seat attachments. If the occupant is allowed to continue to move with little or no

restraint through any significant portion of the

energy-absorbing process anywhere in the system (not

just in the seat and restraint system), a great deal

more stroke or, a much higher load, will be required

to decelerate the occupant. If the occupant moves

with little restriction until the fuselage stops moving, the occupant will then require the same stop- ping distance as the fuselage to experience the same

G loads as the fuselage. Since this stroke is not

available, the loads would be high.

Aside from the seat structure, there are other areas within

the aircraft where energy absorption may find application. Protective padding, generally a plastic foam, should be used

where structure is likely to be impacted by the occupant, par- ticularly where head impact is likely. Deforming structure

such as sheet metal behind the foam also is helpful in such

items as instrument panels, glare screens, etc. Characteristics that should aid in the selection of foams for such appli- cations are discussed in Section 10.9. Also, energy-absorbing webbing for restraint systems and litters is discussed in Section 7.4.4.

4.7.2 Principle of Energy Absorption - Illustration

As an example of the energy-absorption allocations, rewrite

equation (1) for stopping distance as followsi

SNVo02 _ vf 2

s 0 29- (2i

Assuming that v a 42 ft/sec, ve 0, and the average deceleration produced b? deforming terrAin, flattening tires, stroking

energy-absorbing gear, and crushing fuselage is 10 G:

422

(2)(32.2)(l0) " 2.73 ft - 32.87 in.

70
This stroke is 2.73 times the minimum required for the seat;

however, the loads are well within human tolerance limits. If

the entire cumulative stroke could be accomplished et 11.5 G,

which is assumed to produce a deceleration environment tolerable to humans in this direction, the total distance is

422

S (2)(32.2)(il.5) a 2.38 ft a 28.60 in.

Obviously, 28.60 in. of stroke is impractical for a seat, so

the crash energy-absorption function must be a combination of

energy-absorbing landing gear, crushable airframe structure,

and seat energy absorption. The following example illustrates

how the seat and airframe (including the landing gear) combine

to limit decelerative loading of the occupant, assuming rigid

body mechanics, a triangular deceleration input pulse, and a

seat energy absorber load-deflection curve with the same rise

time as the input pulse and a constant limit load.

The triangular deceleration-time plot is an assumed, idealized

input to the system. In actual practice, the dynamic response

of the system as measured on any individual component does not

match this form because of the differing dynamic properties of

the components as discussed in Sectivn 4.7.3.2. The displacement of the seat/occupant system relative to the airframe is computed using the following notation:

Let Gm a maximum airframe deceleration in the vicinity of

mthe seat attachment, G

GL a maximum seat/occupant system deceleration, G

K - GL/Gm, tL/tm (limited to 0.5 or less)

t a time at maximum airframe deceleration (one-half input pulse duration), sec

-'tL - time to reach-maximum system deceleration, sec

t a time, sec

v a velocity, ft/sec

va = velocity of airframe at any time t, ft/sec

vs x velocity of seat/occupant system, ft/sec

.'

71

r*
VL common airframe and system velocity at t t

ft/sec

vo - initial impact velocity, ft/sec

g s acceleration due to gravity, ft/sec2

a - airframe acceleration, ft/sec2

a. = seat/occupant acceleration, ft/sec2

S = displacement, ft

S a - airframe displacement, ft

S8 a seat/occupant system displacement, ft

The airframe acceleration in the interval 0 : t • tm is given

by

a = -Ggt/tm (3)

where the minus sign indicates a deceleration. The velocity

during the same interval, starting from an initial value of vo,

can be found by integration of Equation (3)

va v+ a ÷ adt

avO JGmg(L)dt

Ggt 2

'V0o "2 (4)

The airframe displacement at time t is then

Sa= Vadt

72

I- ____ - ___ - ____ ___ ____ ___ -. - - -- -.------ i . • • .:",[ ,•T ]i , *, T -. -,..,i t. n. u ...l, .n n t
m 2 !m (V• G dt2

- Gmgtm2rtm 2 /6 (5)

For the interval tm < t < 2tM, the airframe acceleration is

a - -Grg + Gmg(t - tm)/tm (6)

and the velocity,

Va nv0 + Iadt

II

t 2

Sv0 + G mg(tm - 2t + 2-t) (7)

30 that, at t = 2 tm

va = v 0 + G g(tm - 4tm + 2 tm)

M V -GMgtm (8)

Since the peak deceleration G is that required to bring the

aircraft to rest at time 2tmI*

Va 0 v0 0- GtmG 

73

--L..-
and

V0 ~G m tM

The airframe displacement at 2t m is then

-' 2

v 0t M tm/6

+ I v + G9( 2t + t )d

m

2v t~ - G ,gt2 (10)

Substituting Equation (9) into Equation (10) the total airframe

di~splacemnent is

S ~VtM *Ggtm2 (

a omm0

The acceleration of the seat/occupant system matches that of

the airframe for 0 5 t i• t ~where t is determined by the

limiting deceleration G 1Iýsing Equations(4an-5,th

veloityanddisplacement of the seat at. t Lcan be-found as

follows:

GmgtL

5e 0 -2t I

74
"",GmgtL

VotL (12)

For t S t t where tf is the time when the seat/occupant systak comes ýdrest f

as a -G-g (13)

and the system velocity in this interval is given by

V GmgtL2 jasdt

Gm tL 2

Vo - 2 t. - GLg(t - tL) (14)

Since v = 0 at t = tf, Equation (14) can be used to find the

final tfme tf

0 V GmgtL2 G

Ou•vO - 2t - GLgtf * GLgtL

Introducing the variable

K - GL/Gm - tL/tm (15)

75

--- -- - - -- -.- -- -- -- .. -.-- -
the time tf can be written

tf 1( K (16)

Using Equation 415) to substitute for t and GL in Equations

(12) and (14), the seat/occupant system %isplacement at t f is

found by

G gK 3t 2 fi K 2G gtm

Ss =VKtm - m + (vo + K 2m m KGmgt)dt

m

F K2 K 32 Kt -2

s mGmg i + 2-)ttf -6

=mgtm + K) 17

The stroke distance required by the seat is the displacement

of Equation (17) less that of the airframe, which is given by

Equation (11):

G -t 2. 1 K K3

Stroke, S = m t m 4 +18) - j-l) 

The above result also can be obtained geometrically, using the

velocity and displacement curves shown in Figure 17. Pcr further clarification, this somewhat simpler procedure is presented below. -

The velocity of the airframe at time t is equal to the initial

velocity plus the change in velocity from t 0 to t -t,

vt -vo + (a). 19 !

£ 0 a 2 (19)

76

-. ______________________ . . . . . - . __________ *:.-, i
GM Input to seat

V ~energy- absorption

system

~Seat/occupant

0 system deceler41 ation

o, GL

ý- t __Oý2t

-Airframe velocityi

>4 Seat/occupant

~system velocity

0 'r1V II

tL 2t L -I

Total seat/occupant - Stroke of seat

4J system displacement -- energy-absorbing

44 system

Airframe displacement

Time Wt, sec

Figure 17. Deceleration-time, velocity-time, and distancetime curves used in analysis of seat/occupant

displacement with respect to the airframe.I

* 77
Substituting the value of a from Equation (3),

S~ fGmgt

vV y + [G~]t t 0 t (2

m1220

V Vt [ m: (20)

Now, assuming that the airframe comes to rest so that v(t - 2 tm) = 0, the total velocity change can be said to equal the

initial velocity. Since this corresponds to the total area

under the deceleration versus time curve,

Vo G mgtm (21)

Substituting Equation (21) into Equation (20) yields

Gg t 22

vt = Ggt -GMg

t =Gmm 2 tm (22)

Using Equation 142), we can now compute the common velocity of

the airframe and the system at time tLI

L Gmgtm 2tm - - 23)

The change in velocity in the time interval tL is '

6v • vo - VL (24)

"i

78 "
Substituting Equations (21) and (23) into Equation (24) yields

GmgtL2 Av Gmgtm - (Gmgtm - 2tmL

2

GmgtL 2t m (25)

The areas of interest in the velocity-versus-time graph in Fig-

* ure 17 can be calculated now using the relationships just derived together with geometrical considerations.

Recognizing that the curve describing the velocity of the airframe consists of the two parabolic segments shown in Figure 18, connected at time t , it can be seen that A is the

area under a parabola of bale t. and height tv. Thhrefore,

AI =• 2 A GmtL2) 2t. / t L -- 3tm gtL2 (26)

VO

4o

0

0

tm 2 tm

Time

Figure 18. Airframe velocity-time curve.

79

-. 4 m -. - •
Area Al is simply a rectangle of base tL and height vL, so

that

(G GigtL2 t G gtL (

ttL gtm /2 GmgtLtm " (27)LtSi 

Since the system is undergoing a constant deceleration beginning at tL, area AV can be represented by the relationship

VL2

2 Av •j i

Substituting from Equation (23) and noting that GL - KGm,

A ( v = n~ G t L 2 ) 2 G g

2KGm-g (28)

The area sought as representing the energy-absorption stroke

of the seat is A In order to solve for this area, A

must first be established. A can be determined by n ig

that, due to the triangular sAie of the acceleration pulse,

the airframe velocity curve consists of two parabolic segments

meeting at the midpoint of the curve, as shown in Figure 18.

If a straight line is constructed joining v and 2t , the two

shaded areas bounded by the curve and the l0e can shown to

be equal since they are both areas between parabolic curves

described by the same basic equation and a secant. The total

area under the curve can then be said to be the same as the

area of the triangle formed by the coordinate axes and line

connecting v0 and 2 tm. Therefore, -

A1 + A1 +A 111 = (.-) 2tm = Votm = gt2 (29)

and

ANV = (AI + AIX AV) - (AI + Ail + AIII) (30)

80
Substituting from Equations (26), (27), (28), and (29) yields

G gtL3 GgtL3

A IV 3t + GmgtLtm 2t

1 4- r2

g G mgt L2 Gmgtm 2

m2gt t 2 2KG g

(GSimplification and substitution of Xtm for tL yields

A =S 2 K3

"I"mm 2 1 (31)

which is the same as Equation (18).

I. As an example, consider a triangular pulse representing a

change in velocity of 42 ft/sec with

G m = 48 G

t = 0.02/ sec

m

K = 1 = 0.24 48

The required stroke is then

0, •024 1 - '0.4)3 I

Stroke = (48)(3•6)(0.027,2 0 4" 2(0.24) -:L + 20.24 244

= 16.25 in.

Test data show this stroke to be less than that required. Much

of this difference can be attributed to system inefficiencies.

It has been found in tests that an efficiency of approximately

t 81

4 . --. ___ . . ____ . -.- ___ -.-- ___ ___
80 percent can be expected from a rod-bending sled decelerator and a wire-bending seat load limiter (References 31 and

32). Therefore, correcting the calculated distance yields

16.25/0.8 - 20.31 in. It must be realized that 20.31 in. is

probably a valid stroke for systems with little or no friction,

such as ceiling-mounted troop seats. For seats guided by sliding or rolling components, friction adds to the resistive force,

thus producing an apparent increase in efficiency. However, in

general, large frictional resistance is not desirable because

of the variation of the net resistive force and hence occupant

decelerative loading as a function of loading direction. Review of the above indicates that the 12-in. minimum seat stroke

required for the design pulse (used in the above calculations)

is hardly adequate and should not be compromised unless other

provisions are included to reduce the residual energy that the

seat is required to absorb.

4.7.3 Dynamic Response

4.7.3.1 Effective Weight: The .cacept of effective weight

has been used to account for ma4jcs supported by components

other than the stroking portion of the seat; e.g., the seat

occupant's lower legs supported by the floor during vertical

loading. The effective weight of the occupant plus the weight

of the movable portion of the seat is multiplied by the limitload factor (G) during calculation of limit loads. The technique is not completely accurate physically because rigid

bodies cannot adequately simulate the dynamic response of the

real system. Seat designs should be dynamically analyzed and

tested to establish their dynamic response and to demonstrate

that they will provide the desired degree of occupant protection.

4.7.3.2 Theoretical System Response: A major design factor

influencing the seat response is the movable seat mass. For

very light seats, the gross response of the occupant can be

estimated using the approximate mass of the occupant acting on

the seat (80 perceizt when considering the vertical direction

as discussed later in this chapter). However, when the seat

31. Reilly, M. J., CRASHWORTHY, TROOP SEAT TESTING PROGRAM,

The Boeing Vertol Company; USAAMRDL Technical Report 4

77-13, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia,

August 1977, AD A048975.

32. Reilly, M. J., CRASHWORTHY, HELICOPTER GUNNER SEAT TESTING

PROGRAM, The Boeing Vertol Company; USARTL Technical Report 78-7, U. S. Army Research and Technology Laboratories,

Fort Eustis, Virginia, February 1978, AD A054970.

82
mass increases to values typical of integrally armored crewseats, interaction between the mass and spring properties of

the seat and occupant can become significant. The occupant S~and seat components then realize sharp deceleration excursions,

• i.e., spikes.

The dynamics of the problem are illustrated in Figure 19, which

presents the theoretical response of an integrally armored crewseat and occupant to an input crash pulse as calculated by a

digital computer analysis (described in Reference 24) and summarized in Section 4.8.6. The analysis simulates the occupant

by three lumped masses representing the head, chest, and pelvis. The cushion and seat are represented by two additional

masses. The five masses are connected by damped springs in

the model.

The response curves for the seat structure, occupant pelvis,

and chest are shown as functions of time for the indicated input excitation. The seat used was an energy-absorbing, integrally armored model set to stroke at 18 G (18 times the effective weight of the occupant plus movable seat). The armored

seat bucket weighed 40.6 lb, and the energy absorber provided

a trapezoidal force-versus-deformation characteristic. It can

be seen that the dynamic response of the seat and segments of

the body are not independent of one another and vary as the

model springs load and unload.

Initially, the seat pan deceleration lags the input pulse as

the springs representing the flesh and the cushion as well as

the elastic spring of the seat structure are loaded. The

stroking force of the energy absorber was sized for a deceleration of a particular mass, and -he effective mass is not yet

being applied to the seat structure because of the incomplete

spring compression. Therefore, the seat pan deceleration exceeds the deceleration required to effect the force necessary

to stroke the energy absorber. The seat pan deceleration approaches 43 G before the cushion and flesh springs compress to

the point that significant deceleration of the pelvis begins.

As deceleration of the pelvic mass increases, an increasing

reaction force is applied in the downward direction ern the seat

pan. Thp seat pan deceleration decreases from A3 G to approximately 27 C as the effective mass is increased.

Because the input decelerative loading is still increasing and

the chest inertial load has not yet been applied to the system,

both the seat pan and the pelvic decelerations increase. As

the spring representing the buttocks flesh and cushion bottoms

out, the pelvic deceleration continues to increase, further

83
50 Initial spike m Input pulse (heavy solid)

\- Vertical seat pan (solid)

-00-Vertical dummy chest (dotted) --.. Vertical dummy pelvis (dashed)

40. ,0 I !

econdary spike

30"

I0

20 0 

U)I / / 0) 0 ,

o I 0 //

10 / 0

//

0 0/

i0I / /

/ /C

/ 0

/Initial notch

-10 Secondary notch

0 26 40 60 80 100

Time, msec

Figure 19. Deceleration versus time for various

components of seat and occupant.

(From Reference 24).

loading the seat pan and decreasing its deceleration. It can

be seen that the Beat pan experiences a small acceleration

under the combined loading of tne occupant's pelvis and chest.

As the chest deceleration increases, the decelerations of the

seat pan and the pelvis tend to normalize near the G level corresponding to the limit-load factor of the energy-absorbing

system.

84
"In summary, the limit load must be set at a load factor considerably below the tolerable level in order to limit the occupant response to a tolerable level, particularly for seats of

high movable mass.

4.7.3.3 Empirical System Response: In the past few years,

several programs were conducted in which crashworthy armored

seats were dynamically tested (References 16, 24, and 33

through 35). These programs included drop tests in which the

seats' response to decelerative loading in the vertical direction was measured. Two types of tests were conducted: in the

first, the impact velocity vector was parallel, but in the opposite direction, to the loading and along the vertical axis of

the seat, the yaw axis related to the aircraft (upward and perpendicular to the aircraft longitudinal axis), and in the second, the seat was pitched forward 30 degrees and rolled 10 degrees relative to the aircraft axis system. These dynamic

tests demonstrated a characteristic deceleration-time history

very similar to that theoretically predicted (see Figure 201.

The characteristic shape has been evident in essentially all

tests to date; however, the magnitudes of the spikes and

notches vary. The characteristic shape of the seat pan

deceleration-versus-time history includes a high initial spike

followed by a deep notch that sometimes passes through zero

and actually becomes an acceleration rather than a deceleration. This notch is followed by a second high spike followed

by various waveforms, damping out and usually centering around

the load factor used in sizing the energy-absorption system

loads.

33. Mazelsky, B., A CRASHWORTHY ARMORED PILOT SEAT FOR HELICOPTERS, ARA Inc.1 USAAVSCOM Technical Report 73-34 and

Report No. NADC-74018-40, Joint Report issued by Naval

Air Systems Command, Washington, D. C., and U. S. Army

Aviation Systems Command, St. Louis, Missouri, January

1974, AD A007551.

34. -Domzalski, L. P. rjand Singley, G. T., III, JOINT APMY/NAVY

TEST PROGRAM FOR UTTAS SEATING SYSTEMS, NADC-79229-60, -

Naval Air Development Center, Warminster, Pennsylvania, to

be published.

35. Dummer, R. J., QUALIFICATION TEST REPORT 613-1787 COOLQUALIFICATION TESTING OF ARMORED CRASHWORTHY AIRCREW SEAT

RA-305 2 5 -1 (FOR SIKORSKY AIRCRAFT CONTRACT 576344), U. S.

Army Contract No. DAAJ01-77-C-0001, Norton Company, Industrial Ceramics Division, Worcester, Massachusetts, revised

January 1979.

85
00

to 0i

0 00

* L)

0) 00 0

>1 0

4)0 0

0 r0

D'U0TIV28TOOQ
IjI

The explanation of the characteristic waveform is associated

• ¶ • with the inherent dynamic response of the seating system and

its occupant. As explained previously, total coupling of the

seat and its occupant is not achieveO since the occupant consists of masses connected by body members, such as the spinal

column and neck, which are not rigid. 4

Further, because the dummy is seated on simulated buttocks

flesh and a comfort cushion, it is not rigidly connected to

the seat pan. Since the energy-absorbing mechanism of the seat

must be set for a given load (calculated by multiplying the

effective weight of the occupant and movable part of the seat

by the desired limit-load factor of 11.5 G), the actual deceleration measured on the seat pan will vary inversely to the

Scoupled weight (ht) according to the relationship a = FP/wt

where a represents the deceleration in G units, F, the load in

pounds resisting the stroke of the seat, and w , the coupled

weight in pounds. The term coupled, as used here, simply indicates that the applicable connecting springs are compressed

sufficiently to result in the body segment being decelerated

in phase at approximately the same rate as the seat pan (as

would a rigidly attached mass).

A deceleration applied to the seat pan initially decelerates

the movable seat mass only. Consequently, deceleration of the

seat pan reaches a large magnitude as indicated by the initial

spike in Figure 20. As the cushion and the simulated flesh on

the buttocks compress, the deceleration of the pelvic mass increases. As the spinal column compresses, the deceleration

of the chest increases. The deceleration of these masses increases as a result of the increased load in the connecting

members. There is a spring constant involved when the connecting members are represented by springs. This constant defines

the relationship between deflection and load; i.e., the further

a spring is compressed, the larger the load required to compress it. The body segments react in the same way. Therefore, the greater the compression, the higher the load, and I

the higher the deceleration of the body segments.

As an illustration, consider Figures 21, 22, and 23. Comparing

the figure of a seat occupant with a system of springs and

masses, it can be seen that when the initial deceleration of

the seat pan commences, the springs in the body are not compressed. The body and seat system have been under a 0-G environment during the drop. Therefore, the springs are totally

unloaded as illustrated in Figure 21. This simply means that

since the springs are not compressed when the deceleration

first commences, large loads cannot immediately be applied to

the body segments. As the pulse continues, the body segments

* 87

- ... --.--- ------.-- . .-.--- i
- 0 -Head

Figure 21. Initial condition, no load.

area is omprespodspingtodclrinla,

a4 Seat

Crash load

Figure 23. Onigh-fdeceleration load wpringn coprelsedc

ara srepodngtodceert88lod

bu4- pertroadhedaent
continue to move undur the resistive load of the partially compressed springs, thus decelerating more slowly than the seat

and building up a velocity relative to the seat pan. Eventually the velocities of the body segments and the seat pan

must all approach a common value. This usually occurs later in

the sequence, after the secondary spike. In the interval, the

deceleration of the seat pan responds as a function of energy

absorber force, input pulse, seat and dummy weight, and spring

and damping characteristics.

Initially, the seet pan deceleration reaches a high value (initial spike). This ccc'"r because the resistive force in the

energy-absorbing systn tq sot at R given value considering

the weight of the movai.'. aortion o' the seat and the occupant.

The seat pan is decelerat.ed initialJy at a magnitude consistent

with the force of the energy-absorAing mechanism divided by

only the weight of the movable part of the seat, which is considerably less than the design weight (the weight of the movable portion of the seat and the effective weight of the occupant). Thus, it is expected that the magnitude of initial seat

pan deceleration will always exceed the limit-load factor for

which the composite energy-absorbing system was designed.

Eventually, the cushion and buttocks springs are compressed,

and the pelvic mass loads into the seat pan (see Figure 22).

The increase of the coupled mass decreases the deceleration of L

the seat pan from its initial peak. The seat pan deceleration

then decreases drastically as evidenced by the initial notch

in the deceleration-time history. At times, when the deceleration actually turns into an acceleration, it simply means that

the mass of the pelvis is receiving a relatively high deceleration and the reaction load is high enough to accelerate the

seat pan towards the aircraft floor. It is apparent that the

magnitude of this notch is a strong function of the spring rate

of the seat. Since the spine normally is still not compressed

significantly, it is not carrying high loads. This is evidenced by the small decelerations measured in the chest, which

is being supported by the spine.

Since this is a dynamically loaded spring system, the springs

asoociated with the buttocks and the cushion can overshoot as

they bottom out during the sequence and then unload again. The

unloading permits the seat pan deceleration to rise again to

the secondary spike on the trace. As the pelvis unloads, the

reaction load on the seat pan decreases and the seat pan deceleration spike can be extremely high. Note that the high deceleration of the seat pan does not necessarily correlate with the

high deceleration of the pelvis or chest. From the data reviewed, both analytical and empirical, it is generally the opposite; i.e., the unloading of the pelvis and/or the chest produces the spike in the seat pan deceleration.

89

il l-IN.i ..11I-ll __________ - . ...
As the cushion and buttocks again load up and the pelvis deceleration increases, the high seat pan deceleration of the

second spike is decreased. Also, the two characteristic deceleration spikes are usually followed by an increased compressive

load in the spine and a buildup of deceleration of the chest.

Eventually, the phasing of the decelerations of the various

system segments begins to converge toward the average load factor for which the limit load of the energy-absorbing system

was designed, as illustrated in Figure 23.

It is informative to note (see Figure 20), that the peak decelerations of the seat pan do not necessarily coincide with peak

decelerations of the human occupant and, thus, are not necessarily hazardous to occupant safety. The Eiband human tolerance data of Section 4.3.3, Volume II, repeated here in Figure 24 for ease of reference, do not present information on the

seat pan deceleration excursions from the average, or uniform

acceleration experienced by the vehicle, and are therefore not

informative on the subject. Again, this is an area presently

being investigated under U. S. Army sponsorship. These criteria will be updated when additional data are available.

The response phenomena described above comprise the predictable

response of the occupant/seat system to the input pulse. The

high decelerations measured on the seat pan are not necessarily

correlated with high decelerations of the occupantl however,

this does not imply that the seat will provide the required protection. The entire deceleration history to which the occupant is exposed must be considered. As pointed out, low decelerations of the seat pan may be accompanied, and caused by,

high loads imposed on the occupant. Thus, it is imperative

that additional information relative to human tolerance to

transient loading in the vertical direction be obtained and

that the criteria for designing vertical energy attenuating

systems for seats be refined and made more comprehensive. Several programs are now underway to expand knowledge in this area.

As the data become available, criteria contained in this guide

will be updated,

4.7.3.4 Tailoring of Energy Absorber: Results of analyses

conducted under a U. S. Navy-sponsored program (Reference 36)

indicated that the force-versus-deformation characteristic of

the energy-absorbing system can be shaped to enable more efficient use of the stroke distance available. As yet, test data

have not substantiated the analyses. However, proper tailoring

36. Carr, R. W., and Phillips, N. S., DEFINITION OF DESIGN

CRITERIA FOR ENERGY ABSORPTION SYSTEMS, Beta Industries

Incorporated; Report No. NADC-AC-7007, Naval Air Development Center, Warminster, Pennsylvania, 11 June 1970,

AD 871040.

90
- -~ -. 4

4-

4J C

1. t41

914
of the waveform should produce the potential for maximization

of the efficiency of a particular system. Research efforts

considering this approach are presently being sponsored by the

U. S. Army. Results should be available by 1981.

4.7.3.5 Adjustable Multiple Limit-Load Devices: Since stroke

in aircraft cockpits, it is extremely important to maximize

the efficient use of the available distance. Energy absorbers

that stroke at a given limit load are sized for the effective

'.,iygit -f the 50th-percentile occupant (by weight). This Implies that the majority of occupants will stroke at or near

the optimum load. Heavier people, however, will stroke at

a lower deceleration level and will require a longer stroke

than the average for which the system was designed, while the

lighter occupant will stroke at a higher deceleration and use

less of the available stroke. Obviously, the energy-absorption

load settings are not optimum for occupants whose weights lie

in the tails of the occupant weight distribution. If adjustment of the limit load for the weight of the specific occupant

is possible, then the system can be optimized for the total

occupant weight distribution.

Variable limit-load energy absorbers can be controlled either

passively (requiring no action by the occupant) or actively

(requiring a conscious action by the occupant). The lassive

device is more complicated and would require a considerably

more sophisticated system to control. It would have to be a

force/time integrating system, since the limit load of the energy absorber could not be a function of the dynamic loading of the seat associated with occupants simply sitting down hard.

There is no doubt that this type of device could be developed; but the cost and weight may be prohibitive, while an actively

controlled device should be neither overly complex nor costly. To achieve most of the advantages offered by the system, the

load would not have to be infinitely adjustable but could be

applied in several increments. The occupant would simply turn

a dial or move a lever to a weight range best fitting his own

weight.

Previous studies (Reference 24) have indicated that a total

excursion of approximately 6 G results-from using a single

limit load set for the 95th-percentile occupant weight. The

6-G excursion could be essentially eliminated with either type

of variable-load energy absorber. This would allow decelerations of all occupants to be nearly identical and enable use

of essentially the same stroke distance for the same decelerative loading and input crash severity for all occupant weights.

92

4.i
Variable limit-load energy absorbers should therefore be incor-.

porated in the vertical direction in all new crashworthy seating systems and retrofit should be considered for seating systems now in use that include stroking capabilities together

with replaceable energy absorbers.

4.8 COMPUTERIZED METHODS OF ANALYSIS

4.8.1 General

Prediction of occupant and seat structure response to dynamic loading presents a complex engineering problem. The use of

computer-aided design in these cases is essential, since the

dynamic interaction of the occupant and the seat/restraint system is much too complex for analysis by manual techniques.

A number of dynamic models of the human body have been developed for use in crash survivability analysis. These models

vary in complexity and possess from I to 40 degrees of freedom

(References 37 through 54). One-dimensional models have been

37. Bacchetti, A. C., and Maltha, J., MADYMO - A GENERAL PURPOSE MATHEMATICAL DYNAMICAL MODEL FOR CRASH VICTIM SIMULATION, Report No. 753012-C, Instituut voor Wegtransportmiddelen, Netherlands 1978.

38. Bartz, J. L., DEVELOPMENT AND VALIDATION OF A COMPUTER

SIMULATION OF A CRASH VICTIM IN THREE DIMENSIONS, Proceedings, Sixteenth Stapp Car Crash Conference, Society

of Automotive Engineers, Inc., New York, 1972, pp. 105-127.

39. Danforth, J. P., and Randall, C. D., MODIFIED ROS OCCUPANT

DYNAMICS SIMULATION USER MANUAL, Publication No. GMR-1254,

General Motors Corporation Research Laboratory, Warren,

Michigan, 1972.

40. Fleck, J. T., Butler, F. E., and Vogel, S. L., AN IMPROVED

THREE DIMENSIONAL COMPUTER SIMULATION OF MOTOR VEHICLE

CRASH VICTIMS, Final Technical Report No. ZO-5180-L-l (in

four volumes), Calspan Corporation, Buffalo, New York, - 1974.

41. Furosho, H., Yokoya, K., and Fujiki, S., ANALYSIS OF OCCUPANT MOVEMENTS IN REAR-END COLLISION, Paper No. 13, in

Safety Research Tour in the U.S.A. from the Viewpoint of

Vehicle Dynamics, 1969.

42. Furosho, H., and Yokoya, K., ANALYSIS OF OCCUPANT'S MOVEMENT IN HEAD-ON COLLISION, Transactions of the Society of

Automotive Engineers of Japan, No. 1, Tokyo, Japan, 1970,

Spp. 

145-155.
used in prediction of human body response to an ejection seat

firing (References 55 through 57), which, if the body is tightly

restrained, can be approximated as a one-dimensional phenomenon.

However, a vehicle crash generally involves a horizontal component of deceleration, which forces rotation of body segments

with respect to each other. If no lateral component of deceleration is present, a two-dimensional model will suffice, provided the restraint system is symmetrical. However, lateral

43. Glancy, J. J., and Larsen, S. E., USERS GUIDE FOR PROGRAM

SIMULA, Report TDR No. 72-23, Dynamic Science, Division of Ultrasystems, Inc., Phoenix, Arizona, 1972.

44. Huston, R. L., Hessel, R., and Passerello, C., A THREEDIMENSIONAL VEHICLE-MAN MODEL FOR COLLISION AND HIGH ACCELERATION STUDIES, Paper No. 740275, presented at Automobile Engineering Conference, Society of Automotive

Engineers, Inc., Detroit, Michigan, 25 February - 1 March

1974.

45. McHenxy, R. R., ANALYSIS OF THE DYNAMICS OF AUTOMOBILE

PASSENGER-RESTRAINT SYSTEMS, Proceedings, Seventh Stapp

Car Conference, Society of Automotive Engineers, Inc.,

New York, 1963, pp. 207-249.

46. Robbins, D. H., THREE-DIMENSIONAL SIMULATION OF ADVANCED

AUTOMOTIVE RESTRAINT SYSTEMS, Paper No. 700421, In 1970

International Automotive Safety Conference Compendium

P-30, Society of Automotive Engineers, Inc., New York,

97 0.

47. Robbins, D. H., Bennett, R. 0., Jr., and Bowman, B. M.,

USER-ORIENTED MATHEMATICAL CRASH VICTIM SIMULATOR, Proceedings, Sixteenth Stapp Car Crash Conference, Society of

Automotive Engineers, Inc., New York, 1972, pp. 128-148.

48. Robbins, D. H., Bennett, R. 0., and Roberts, R. L., HSRI

TWO-DIMENSIONAL CRASH VICTIM SIMULATOR: ANALYSIS, VERIFICATION, AND USER'S MANUAL, Final Report, Report No.

HSRI-Bio-M-70-8, Highway Safety Research Institute, University of Michigan, Ann Arbor, Michigan, 1970.

49. Robbins, D. H., Bowman, B. M., and Bennett, R. 0., THE

MVMA TWO-DIMENSIONAL CRASH VICTIM SIMULATIONS, Proceedings, Eighteenth Stapp Car Crash Conference, Society of

Automotive Engineers, Inc., New York, 1974, pp. 657-678.

50. Segal, D. J., REVISED COMPUTER SIMULATION OF THE AUTOMOBILE CRASH VICTIM, Report No. VJ-2759-V-2, Cornell Aeronautical Laboratory,, Inc., Buffalo, New York, 1971.

94
loading is common in helicopter accidents. Also, the diagonal

S• shoulder belt used in some troop restraints is asymmetrical

and may cause lateral motion.of the occupant even in the absence of a lateral deceleration. Therefore, a model that is

generally useful in restraint system evaluation must be capable of predicting three-dimensional motion, and several threedimensional kinematic models, made up of interconnected rigid links, have been developed (References 38, 44, 47 and 53).

Subsequent sections of this chapter describe the models of possible use to designers of seats and restraint systems.

4.8.2 Program SOM-LA

In 1972, the FAA initiated a program to provide a practical

engineering tool for use in the design and evaluation of seats

and restraint systems for light aircraft. This program incorporated a dynamic model of the human body combined with a

51. Segal, 0. J., and McHenry, R. R., COMPUTER SIMULATION OF

AUTOMOBILE CRASH VICTIM - REVISION, Report No. VJ-2492-I,

Cornell Aeronautical Laboratory, Inc., Buffalo, New York,

1967.

52. Twigg, D. W., and Karnes, R. N., PROMETHEUS, A USERORIENTED PROGRAM FOR HUMAN CRASH DYNAMICS, Boeing Computer

Services, Report No. BCS 40038, Department of the Navy,

Office of Naval Research, Waahington, D. C., 1974.

53. Young, R. D., THREE-DIMENSIONAL MATHEMATICAL MODEL OF AN

AUTOMOBILS PASSENGER, Research Report 140-2, Texas TransSportation institute, College Station, Texas, 1970.

54. Young, R. D., Ross, H. E., and Lammert, W. F., SIMULATION

OF THE PEDESTRIAN DURING VEHICLE IMPACT, Paper No. 27,

Proceedings, Third International Congress on Automotive Safety, Vol. II, Society of Automotive Engineers, Inc., New York, 1974.

55. Kroeger, W. J., INTERNAL VIBRATIONS EXCITED IN THE OPERATION OF PERSONNEL EMERGENCY ESCAPE CATAPULTS, Memorandum

Report 340, Frankfort Arsenal Laboratory Division, Philadelphia, Pennsylvania, 1946.

56. Latham, W. F., A STT')Y IN BODY BALLISTICS: SEAT EJECTION, Proceedings of Royal Society, London, England, 1957,

B 147: 121-139.

57. Stech, E. L., and Payne, P. R., DYNAMIC MODELS OF THE HUMAN BODY, Frost Engineering Development Corp., AMRL Technical Report 66-157, Aerospace Medical Research Laboratory,

Wright-Patterson Air Force Base, Ohio, 1969.

95
finite element model of the seat structure. The program is

intended to provide the designer with a tool with which he

can analyze the structural elements of the seat as well as

evaluate the dynamic response of the occupant during a crash.

The digital computer program based on this model is called

SOM-LA (Seat/Occupant Model: Light Aircraft).

The aircraft occupant is modeled by twelve rigid mass segments with rotational springs and dampers at the joints. The segments represent the head, neck, upper torso, lower torso, upper

arms, forearms, thighs, and lower legs, as shown in Figure 25.

Each of the torso joints possesses three rotational degrees of

freedom. The elbows, the knees, and the head-neck joint are

hinge-type connections, each of which is allowed one additional

degree of freedom. In total, the occupant model possesses 29

degrees of freedom.

Rotations at the body joints are resisted by torsional springs and dampers, whose characteristics depend on user selection of

human or dummy occupant.

External forces are applied to the body segments by the seat

cushions, the floor, the belt restraint system, and an optional inflatable restraint. The four available restraint system configurations consist of a lap belt alone or combined with a

single diagonal belt over either shoulder, or a double shoulder belt. A lap belt tiedown strap may be used with the double I shoulder belt system. The restraint loads are transmitted to

the occupant through ellipsoidal surfaces to the upper and

Slower torso segments; and, the points of application depend on

current belt geometry. The capability of the belts to move

relative to the torso surfaces allows simulation of submarining

under the lap belt as well as prediction of the lateral motion

which may result with a single diagonal shoulder belt.

For calculation of external forces exerted on the occupant by

the seat cushions and restraint system, and for prediction of

impact between the occupant and the aircraft interior, 24 surfaces are defined on the body. These surfaces are ellipsoids,

spheres,- and cylinders, as shown in Figure 26. .

The user may select either a finite element model of the seat

structure or a rigid seat representation. The latter can include nonlinear translational and rotational stiffness elements

for simulation of an energy-absorbing seat as shown in Figure 27. The bucket is assumed rigid, and frame elasticity is

modeled by the torsional spring.

The program has been run on CDC 6600, Univac 1108, and IBM 370

systems. Input data include force-deflection information for

96

B.-7
j R#$. 

4 

• • -. . e -- • ••..... vI12

SSegment mass center

e25. Twel -egm t occupant model.

97
1

20 4 2 41 7 1 13 11

Figure 26. Occupant . ; KI model contact surfaces.

I

KT: Translational stiffness (energy absorber)

KRIt: Rotational stittngg8

K T

Rigid

frame 

R Rigid bucket

Figure 27. SOM-LA energy-absorbing seat model.

98
the cushions and belts; crash conditions, in terms of intial

velocity and attitude and time variations of six acceleration

components; occupant description; seat design data; and, if

the prediction of impact with the aircraft interior is desired,

a description of the cabin surfaces. Output data include time

histories of occupant segment positions, velocities, and accelerations; restraint system loads; seat deflections and

forces; details of contact between the occupant and the aircraft interior; and several measures of injury severity. The

injury criteria used in the program are all computed from segment accelerations. The dynamic response index (DRI) provides

an indication of the probability of spinal injury due to a vertical acceleration parallel to the spine. The Severity Index

is calculated for the chest and head, and the Head injury Criterion of Federal Motor Vehicle Safety Standard 208 is also

computed.

This original model is described in Reference 58. Since its

publication, a number of modifications to the model have been

made to improve simulation quality and to provide increased

capability and additional desirable output. In that interim,

an extensive testing program was initiated by the FAA Civil

Aeromedical Institute (CAMI) to provide data for validation of

the model. Work on model improvement and validation is continuing, as described most recently in Reference 59.

4.8.3 Calspan Corporation - CVS

Probably the most sophisticated biomechanical model of the human body intended for crash simulation is the Calspan Corporation Crash Victim Simulator (CVS). Originally reported on in

1972 (Reference 38), the program includes a body dynamics model

with 40 degress of freedom and a contact model that generates

forces from contact with vehicle surfaces. The extensive validation effort has included the following experiments:

58.- Laananen, D. H., DEVELOPMENT-OF A SCIENTIFIC BASIS FOR

ANALYSIS OF AIRCRAFT SEATING SYSTEMS, Dynamic Science,

SDivision of Ultrasystems, Inc.; Report No. FAA-RD-74-130,

U. S. Department of Transportation, Federal Aviation Administration, Washington, D. C., 1975, AD A004306.

59. Chandler, R. F., and Laananen, D. H., SEAT/OCCUPANT CRASH

DYNAMIC ANALYSIS VALIDATION TEST PROGRAM, Paper No. 790590,

presented at Business Aircraft Meeting, Society of Automotive Engineers, Inc., Wichita, Kansas, April 1979.

99

- -. *.... . .. • t .. a-.--.
* Static bench tests with a spherical membrane and

spherical contact surfaces to validate the air bag

shape and contact force algorithm.

e Pendulum tests with a dummy torso form restrained

and decelerated with an air bag to further validate

* Tests with instrumented anthropomorphic dummies on

an impact sled at 20 and 30 mi/h with both belt and

air bag restraints, in which both planar and nonplanar dummy responses were produced.

* A head-on, laterally offset, car-to-car crash test,

with the primary vehicle containing two instrumented

anthropomorphic dummies.

A graphics display model provides rather sophisticated threedimensional views of occupant response, as shown in Figure 28.

Present capabilities of the program, a user manual, and a de-

"scription of its validation are presented in Reference 40.

Figure 28. CVS graphics display model.

4.8.4 PROMETHEUS

In 1572, Boeing Computer Services began work on modification of

a two-dimensional occupant model called SIMULA, which had been

developed earlier by Dynamic Science, Inc. and Arizona State

University. Their final product, which includes interactive,

user-oriented capabilities, is called PROMETHEUS (Reference 52).

100

- - - ' .- -- -- --- -• - "k "
W_'

PROMETHEUS simulates a crash victim with either a twodimensional, seven-link, side-facing mathematical model, shown

in Figure 29(a), restrained by a seat belt and shoulder harness, or an eleven-link, forward-facing, unrestrained model,

shown in Figure 29(b). A nonlinear finite element model of

the impacting structure is incorporated. A new, fast differential equation solver was developed for the program to efficiently compute the transient response of the finite element

vehicle structure and rigid-link occupant in a crash situation. The program is an interactive, user-controlled system

designed for the rapid analysis/data edit/reanalysis cycles

necessary for efficient parametric studies. PROMETHEUS input

aids include free-field data input and an on-line data edit

capability. Output provides user-selected time history and

occupant configuration plots, as well as abbreviated output

lists for rapid scan of results. The program operates on the

CDC 6600 computer in either a batch or an interactive mode.

;Joint 9 Joint

Point mass IDistributed mass

(a) Side-facing model (b) Forward-facinq model.

Figure 29. PROMETHEUS occupant model.

101
a.-n.n n- n I I I I I l l l ll4.8.5 Air Force Head-Spine Model

Under the sponsorship of the U. S. Air Force Aeromedical Re- I search Laboratory, a three-dimensional, discrete model of the

* human spine, torso, and head was developed for the purpose of

evaluating mechanical response in pilot ejection. It was developed in sufficient generality to be applicable to other

body response problems, such as occupant response in aircraft

crash and arbitrary loads on the head-spine system. There are

no restrictions on the distribution of direction of applied

loads, so a wide variety of situations can be treated. The

model is described in Reference 60.

The anatomy is modeled by a collection of rigid bodies, which

represent skeletal segments such as the vertebrae, pelvis,

head, and ribs, interconnected by deformable elements, which

represent ligaments, cartilageneous joints, viscera, and connective tissues. Techniques for representing other aspects of

the ejection environment, such as harnesses and the seat geometry, are included. The model is valid for large displacements

of the spine and treats material nonlinearities. The elements

of the model are illustrated in Figure 30.

The basic model is modular in format, so that components may

be omitted or replaced by simplified representations. Thus,

while the complete model is rather complex and involves substantial computational effort, various simplified models that

are quite effective in duplicating the response of the complete model within a range of conditions are available. Three

methods of solution are available for the analysis: direct

integration in time by either an explicit, central difference

method; by an implicit, trapezoidal method; or by a frequency

analysis method.

A variety of conditions have been simulated, including different rates of onset, ejection at angles, effects of lumbar curvature, and eccentric head loadings. It has been shown that

large initial curvatures and perfectly vertical acceleration

loadings result in substantial flexural resporse of the spine,

which cause large bending moments. It has been further shown

that the combination of the spine's low flexural stiffness,

initial curvature, and mass eccentricity are such that stability cannot be maintained in a 10-G ejection without restraints

or spine-torso-musculature interaction.

60. Belytschko, T., Schirver, L., and Schultz, A., A MODEL FOR

ANALYTIC INVESTIGATION OF THREE-DIMENSIONAL HEAD-SPINE

DYNAMICS 

- FINAL REPORT, University of Illinois at Chicago

Circle; AMRL Technical Report 76-10, Aerospace Medical

Research Laboratory, Wright-Patterson Air Force Base,

Ohio, April 1976, AD A025911.

102
Figure 30. Three-dimensional head-spine model. (From Reference 60)

The the complete rib cage 

models and viscera were used 

on mainly to study the effects of spinal response. The flexural stiffness of the torso is increased substantially by a visceral model, 

addition, 

even though it has no inherent flexural stiffness. In

axial loads.the viscera provide significant reductions in the

4.8.6 One-Dimensional Seat/Occupant Models

Although a three-dimensional simulation should be used for com- plete prediction of aircraft occupant dynamics in investigating

"4•. 103
restraint system properties or cockpit configurations to eliminate secondary impact hazards, the more simple one-dimensional

models also may be useful in crashworthy seat analysis. For

example, a model such as that illustrated in Figure 31 provides

an economical means of optimizing energy absorber characteristics, which would be simulated by spring X . Energy absorber

force-deflection characteristics might be ýkried while searching for the most favorable occupant response, evidenced by a

minimum of spinal deflection, head acceleration, etc. The most

notable difficulty with the use of such a model lies in obtaining valid occupant properties, i.e., masses and spring characteristics. One such model that has been used in seat evaluation is described in Reference 61.

Another widely known one-dimensional model is used to compute

the Dynamic Response Index (DRI). The DRI is a predictor of

spinal injury due to ÷G acceleration, and is based on the response of a single-degree-of-freedom model as described in detail in Volume II.

61. Auyer, W., and Turnbow, J., A STUDY OF THE DYNAMIC RESPONSE OF A DAMPED, MULTI-DEGREE OF FREEDOM, SPRING-MASS

SYSTEM WHICH SIMULATES A SEAT, SEAT CUSHION, AND SEAT OCCUPANT SUBJECTED TO A VERTICAL IMPACT ACCELERATION, Aviation Safety Engineering and Research (AvSER), Division of

Flight Safety Foundation, Inc. (unpublished report).

104
xs

He ad

M 5

Spna

Nck C4 ___

sea Cushion, anocpat

I 105 M

*lo

_torso

105

, .i ,'. -' • • " ,' •,. 'K . 4

[L••,. ~~~~pia ... .. .... K.y. .... n
. 5. ENERGY-ABSORBING DEVICES

5.1 INTRODUCTION

A multitude of devices for absorbing energy and limiting loads

have been proposed, developed, and tested. As demonstrated

earlier, the kinetic energy of a moving mass can be absorbed

by applying a force over a distance; this is the primary mechanism for absorbing crash energy. The larger the distance

through which the force acts, the lower the average load on

the mass. Energy-absorbing mechanisms in aircraft structures

which transmit crash forces to the occupant should stroke at

loads tolerable to humans and should provide stroke distances consistent with these loads and with the energy to be absorbed.

Past experience has shown that plastic deformation of material,

primarily metal, results in a reasonably efficient energy- absorbing process. Consequently, most load-limiting or energy- absorbing devices use that principle.

Desirable features of energy absorbers are as follows:

* The device should provide a predictable force-versusdeformation trace.

* The rapid loading rate expected in crashes should

not cause unexpected changes in the force-versusdeformation characteristic of the device.

* The assembly in which the device is used should have

the ability to sustain tension and compression.

(This might be provided by one or more energy absorbers, or by the basic structure itself, depending

on the system design.)

"* The device should be as light and small as possible.

"* The specific energy absorption (SEA) should be high.

* Thc device should be economical.

"* The device should be capable of being relied upon to

perform satisfactorily throughout the life of the

aircraft (a minimum of 10 years or 8000 flight hours)

without requiring maintenance.

"* The device should not be affected by vibration, dust,

dirt, or other environmental effects. It should be

protected from corrosion.

106
e The device(s) should decelerate the occupant in the

most efficient manner possible while maintaining the

loading environment within the limits of human tolerance.

The discussion that follows refers to load limiters as separate devices. This is not meant to imply that load limiters

must be separable devices at the exclusion of the integral design concept wherein the structure itself is designed to collapse in a controlled and predictable fashion. Rather, the

discussion is present*1 in this way to simplify portrayal of

different methods of absorbing energy and limiting loads.

Research on simple, compact, load-limiting devices has been

conducted by the Government and by private industry. These

data are recorded in References 62 through 70. A brief discussion of some of the more common energy-absorption devices and

concepts applicable to seats is presented in the following text

and in Table 3.

62. Ezra, A., and Fay, R. J., AN ASSESSMENT OF ENERGY ABSORBING DEVICES FOR PROSPECTIVE USE IN AIRCRAFT IMPACT SITUATIONS, in Dynamic Response of Structures, G. Herrmann and

N. Perrone, eds., Pergammon Press, Elmsford, New York,

1972, pp. 225-246.

63. Reilly, M. J., CRASHWORTHY TROO" S.ýAT INVESTIGATION, The

Boeing Vertol Company; USAAMRDL Technical Report 74-93,

Eustis Directorate, U. S. Army Air Mobility Research and

Development Laboratory, Fort Eustis, Virginia, December 1974, AD/A-007090.

64. Kroell, C. K., A SIMPLE, EFFICIENT, ONE SHOT ENERGY ABSORBER, Reprint from Bulletin No. 30, Shock, Vibration,

and Associated Environments, Part III, General Motors Research Laboratory, Warren, Michigan, February 1962.

65. Guist, L. R., and Marble, D. P., PREDICTION OF THE INVERSION LOAD OF A CIRCULAR TUBE, NASA Technical Note D-3622,

Ames Research Center, Moffett Field, California, June 16,

1966.

66. Haley, J. L-, Klemme, R. E., and Turnbow, J. W., TEST AND

EVALUATION OF 1000-4000 POUND LOAD-LIMITING DEVICES, Dynamic Science, AvSER Facility Report M69-2, for U.S. Army

Aviation Materiel Laboratories, Fort Eustis, Virginia,

February 1969.

107

i. . ...
0 I -o 4.6 4.4 a,4 4 a

04. 60.

C4 0

31460 4 0

4.A 4.O j

4.~ C1 ' 0

C 9 9 -06 U

4,49 - a6 4

-- -~'0 - - - .91

'4" 15 x 0 7

6. 40 4L w0 a. -

C CC C-o 0. 96

~' 0 ' 50 4.1 0 41

o w a o9 - 0 9 ,

a , 96 C ~ . 6.

cc c

go-

-4. 4 - :. 

9~~~~ N )9C r 

'4~ '0 ~ v A U c O

'4~~~~ '0:99U 0. v .4

A" I. ft9

So. V . U 9-

c X,4~

9l1 4P c .I, cV AoU 41 9. m 6 V . 2

00 0 mu O

41~9. ACA 1

-'4C CAw 0 *1 . c6 >1 4 '0 C C #.44 40- 640. # 0 #49 o 0 I-C 

> o4. 0 9. C

C 9 4 0 5 av 0 .06 M*A V 4'

C C -969 A -108
In Table 3 Long-Term Reliability refers to the ability of the

device to perform its function without benefit of maintenance

throughout the life of the aircraft. The weight used in calculating SEA values includes the necessary end fittings required

to apply the load except as noted.

Pertinent characteristics of each device listed in Table 3 are

discussed in succeeding paragraphs. The concepts that have

found use in actual seat designs are presented first.

5.2 WIRE OR STRAP BENDING

This device uses the force required to bend a metal wire or

strap around a die or roller. It can be as simple as a steel

wire threaded through a perforated plate or a wire wound around

rollers. One characteristic that may be a problem with this

device (as with all devices affected by or utilizing friction

from metal-to-metal contact) is that an initial peak load

higher than the normal stroking load is induced. This initial

load increase can be reduced or eliminated by providing initial

slack in the wire when passing it over the rollers. These devices, by themselves, do not have the ability to sustain compressive loads. However, by anchoring both ends of the wire

and attaching the seat bucket to the rollers, compressive as

well as tensile loads can be sustained.

67. Rich, M. J., VULNERABILITY AND CRASHWORTHINESS IN THE DESIGN OF ROTARY-WING VEHICLE STRUCTURES, Paper No. 680673,

presented at Aeronautic and Space Engineering and Manufacturing Meeting at Los Angeles, Society of Automotive Engineers, Inc., New York, October 1968.

68. Bendix Products Aerospace Division, ENERGY ABSORBING CHARACTERISTICS OF CRUSHABLE ALUMINUM STRUCTURES IN A SPACE

ENVIRONMENT, Report No. SPP-65-107 (NASA-CR-65096), prepared for NASA Manned Spacecraft Center, Houston, Texas,

July 1965.

69. McGehee, J. R., A PRELIMINARY- EXPERIMENTAL INVESTIGATION

OF AN ENERGY-ABSORPTION PROCESS EMPLOYING FRANGIBLE METAL

TUBING, NASA Technical Note D-1477, National Aeronautics

and Space Administration, Washington, D. C., October 1962.

70. Schwartz, M., DYNAMIC TESTING OF ENERGY-ATTENUATING DEVICES, NADC Report No. AC-6905, Naval Air Development

Center, Warminster, Pennsylvania, October 1969.

)- 109

I-. - -
'.

Two variations of the wire-bending device have been developed

and used in the ceiling- and floor-mounted troop seat illustrated in Figure 32. The two tension-type devices at the top I

of the troop seat are shown in greater detail in Figure 33.

i

Aic Aft rcriaft flo IJ

I

11

I. i

I

Figure 32. Crashworthy troop seat. (From Reference 63)

In the analysis of energy absorbers for the troop seat, re-I

ported in Reference 63, wire of varying diameter was investi- gated in order to produce a notched force-deflection curve as

recommended in Reference 36. It was concluded that the notched

110I

" ' • .. .... ... - -,,-_-- - - ' • I ;, hi'"
Support point Support point

Music wire

Rollers

Top Housing of

j '- /seat

Support pir. back

Sre Stroke ..... _ "_ •mn12 in.

mi~n.

op of

sect back

Seat back

Normal Pully stroked

Position Position

1600

1400

1200

1000

Design load 1,020 lb

800

600

400

200 (1) Initial vertical attenuator (heavier wire)

(2) Revised vertical attenuator (0.100-in. diameter wire)

0

0 1 2 3 4 5 6 7 $ 9 10 fll- 12 13 14

Deflection, in.

Figure 33. Troop seat tension energy absorber includ4 ng

characteristics for two wire diameters.

t -111

SC
.1

force-deflection curve was not suitable for lightweight troop *

seats due to the sensitivity of the system response to location

of the notch in the load-versus-deflection characteristic. A

fixed location for the notch was not compatible with the various dynamic response phasing resulting from the wide range of

troop and equipment weights. The trapezoidal force-deflection

curve produced by the constant limit-load device, although not

as efficient theoretically and ideally as the notched curve

for a specific dynamic condition, appeared to be more tolerant

of the wide range of seat occupant weights. Figure 33 shows

the force-deflection characteristics of the device that were

measured for two different wire diameters. I

I

The other variation of the wire-bending energy absorber used

in the above mentioned troop seat, and shown in Figure 34, is

capable of functioning in tension or compression. The deviceu

is contained in two telescoping aluminum tubes. A cap is

placed on the inner end of the inner tube. Music wire of

0.100-in. diameter, in the shape of a hairpin, is looped i

through the cap, and the two free ends are secured to a stud

in the outer end of the inner tube. A trolley consisting of three rollers sandwiched between two plates bends the wire as' the trolley moves back or forth on the wire. The trolley is

pinned to the outer tube, and a slot is provided in the inner

tube to allow passage of the pin connecting the two.

5.3 INVERSION TUBE

This device uses the force required to invert (to turn insideout or outside-in) a length of metal tubing. The concept

was developed by an American automobile manufacturing company for incorporation into steering columns to produce controlled collapse loads (see Reference 64). No real disadvantages have been noted in experimental tests to date except with

?hose loaded in compression. In dynamic tests of troop seats

(Reference 71) using these devices in compression, there was a

tendency for the outer and inner tubes to misalign, which resulted in failure and crippling of the inner tube. However,

this problem can be solved by using an internal guide to keepthe initial eccentricity from developi-ng. It is possible that

atmospheric corrosion could occur in the closed space between I

the inner and outer tube walls, especially in the bend radius.

It has been suggested that this potential problem might be

solved by injecting a low-density, closed-cell plastic foam

into the small volume between the inner and outer tube walls

71. Singley, G. T., III, FULL SCALE CRASH TESTING OF A CH-47C

HELICOPTER, paper presented at 32nd Annual National V/STOL

Forum, American Helicopter Society, Washington, D. C.,

May 1976.

6I

112

-------------*.-..
•Music wire

Outer alumiunum tube

TInner aluminum tube

1600*r Slot in inner tube

"'"°°

o4

1600

400'-80 ; •

-eslcticn, in.

Figure 34. Tubular 

force-deflection strut wire-bending energy absorber with c orves. (From Reference 31)

113 5

4 V......- - -. - .. - r 4 . -.- - -•:n•- ----- -.-. '. -lk- • . ... ' - -. ". . . .- c -'_•' -. .
to prevent moisture penetration of this area. Also, the tubes

could be plated and/or coated to protect them from corrosion.

The materials used so far in inversion tubes have been 3003-H14

aluminum and mild steel, as described in References 64 through

66. It is possible that an annealed, higher strength alloy

steel, such as 4130 or stainless steel, could yield even higher

specific energy values than those shown in Table 3. However,

the aluminum devices that are in use are both compact and

lightweight.

Figure 35 illustrates a specific design concept of the inversion tube energy absorbers (Reference 72). The load-deflection

curves are also shown for nine devices, three of which were

subjected to fatigue testing and three to both environmental

and fatigue testing. The traces are essentially flat for the

entire stroke distance after the initial peak.

Typical load-deformation information for tubes made from

3003-H14 aluminum is presented in Figure 36. These curves show

the variation in the load-deformation curve for a total of 12

dynamic test specimens and 2 static test specimens. Note that

the peak dynamic load is about 1300 lb while the minimum dynamic load is about 1010 lb. This 25-percent variation is

higher than the variation now being experienced at equivalant

loading rates in operational systems. Posttest examination of

the specimens showing 15- to 25-percent load increases indicated that the inverting tube had been scrubbing the inner

wall of the loading tube due to asymmetric loading in the jig;

therefore, friction probably caused some of the load increase.

Figure 36 shows clearly that the inverting load remains essentially constant during the stroke (Reference 64).

5.4 ROLLING TORUS

Early versions of this energy absorber consisted of a number

of torus elements located in the annular space between two

telescoping cylinders. Modification of this concept has re--

sulted in the substitution of a continuous helix of stainless

steel wire for the toroidal elements. The interference fit

between the cylinders and tori, or wire, causes the wire to

roll when axial loads are applied. The cyclic plastic deformation of the rolling tori or *ire helix and elastic deformation of the tubes effect the energy absorption. The cylinders

remain intact and do not plastically deform when subjected to

S7 esjardins, S. P., et al., CRASHWORTHY ARMORED CREWSEAT

FOR THE UH-60A BLACX HAWK, paper presented at 35th Annual

National Forum, American Helicopter Society, Washington,

D. C., May 1979.

114
RMW

* Thin-walled aluminum

roll tube (3003-H14)

Unstroked

Outer tube_

Partially stroked Plasticall

deformingA

region I

1500j

100

o toThree devices previously en- Three new

Three devices vironmentally devices used

previously and fatigue as control fatigue tested tested group

50 --- S/N 1002 ---- &S/N 1005 u---S/N 1008

.......... SIN 1003 S/N 1006 ~- SIN 1009 ......... S/N 1004 -- uS/N 1007 - S/ IN 1010

0

0 2 4 6 8 10 12 14 16 18 20I

Stroke, in.

Figure 35. Inversion tube concept with typical

force-deflection characteristic.

* 115
6061T-6, 0.035-in.

wall X 2.125-in. 10.25

OD loading tube

d 5.5 n- Stroke

2.0 in.- 3003-H14, 0.035-in.

wall X 1.75-in. OD

inverting tube

2000

Wt 0.35 lb each Load discontinued

1500 iooal : '.. -

1000 • -- ' • .. - •' C d-.. . • • •

S 500 --- Max dynamic Results based

Min dynamic on 2 static and

- Static 12 dynamic tests 0 1 2 3 4 5 6

Deformation (compression), in.

Figure 36. Comparison of dynamic and static loaddeformation curves for inversion tubes.

impact loading. The impact force is transmitted through the

tubes to the tori or wire helix. Dynamic testing of these devices is reported in Reference 70.

The load !imiters using wire as the working medium (Figure 37)

are normally made with cylinders that range from 1 to 2 in. in

diameter with a wall thickness of approximately 0.035 in. The

wire ranges between 0.030 and 0.035 in. in diameter and is of

300 series stainless steel. These bidirectional devices may beused repeatedly several times until fatigbe failure of the wire

occurs. The recent investigation of a lighter weight aluminum

energy absorber of this type is documented in Reference 73.

73. Mazelsky, B., INVESTIGATION OF AN ALUMINUM ROLLING HELIX

CRASH ENERGY ABSORBER, ARA, Inc.1 USAAMRDL Technical Report 77-8, Eustis Directorate, U. S. Army Air Mobility

Research and Development Laboratory, Fort Eustis, Virginia, May 1977, AD A042084.

116

W'_ I
i. 4000 "

3000"

2000-r MI

01 2 3 4

Deflection, in.

Section

Outside

Figure 37. Rolling torus energy absorber.

4.11

• 117

......-,,..
Devices of this type can be single or multiple staged. The

multiple-staged energy absorbers include three tubes with

helices of wire between the walls of the outer tube and the

center tube, and between the center tube and the inner tube.

In operation, one helix of wire is rolled to the end of its

stroke and then the second stage is initiated and rolled.

Staged energy absorbers provide increased stroke distance

without an appreciable increase in prestroked envelope.

The device produces a somewhat jagged load-versus-deformation

characteristic as can be seen in Figure 37. Further, the interference contact between the tori and the cylinders, the

closed spaces between the tube walls, and the spaces between

the wire wraps are prime areas for corrosion. This potential

should be considered during the development, test, and usage

of this device. The long-term effects on performance of the

interference fit between the wire and the tubes is another area

for concern. J

Seats with energy-absorbing mechanisms utilizing this device

are now in use in a modified U. S. Marine helicopter (Reference 16) and in a utility helicopter developed for Iran.

5.5 CRUSHING HONEYCOMB

This device uses the force required to crush or deform a column

of low-density material. In order to provide sufficient column

stability and transverse load resistance, it appears that most

applications will require a telescoping cover to give a d4itional bending strength. Table 3 shows this device to be above

average in all categories with the exception of rebound load-.

ability. Rebound load capacity could probably be added by the incorporation of a suitable mechanism that allows movement in

only one direction.

This device, besides being used on seats, is used as a load

limiter in the main landing gears of some helicopters. In

these applications, the crushable material is installed above

the oleo piston as outlined in Reference 67. The energyabsorption ability of these devices has been responsible for

preventing major structural damage to several aircraft in severe accidents.

To date, the best crushable material for use in this type of

device appears to be corrugated aluminum foil backed by flat

foil, cemented at the nodal points as illustrated in Figure 38. Further research information on the development of

crushable aluminum columns may be found in Reference 68.

118

i • !im i • • ld "....i...............i • •-............. .- "-
Load

S~Loadl

Cement points

Start of

wrapping

Figure 38. Illustration of corrugated aluminum

foil formed into annular column.

5.6 EXTENSION OF BASIC METAL TUBE OR FLAT STRAP

This concept uses the inherent plasticity of certain ductile

metals which elongate under a relatively constant force. The

primary problem with this device is strain concentration at

the end connections. Research to date indicates that annealed

stainless steel in the AISI 300 series is least susceptible to

strain concentrations because of its excellent ductility (45 to

50 percent).

The flat strap device was evaluated for use as a vertical load

limiter for a pilot's seat by the U. S. Naval Aircraft and Crew

Systems Technology Directorate, now part of the U. S. Naval

Air Development Center and was found to perform satisfactorily

(Reference 74). Since a flat strap sustains only minimum compressive loads, a separate rebound device would be necessary

for application in personnel seats-.-

The thin-walled tube will perform in much the same manner as

the flat strap, and it has the advantage of sustaining higher

74. Woodward, C. C., et al., INVESTIGATION, DESIGN AND DEVELOPMENT OF AN F7U-3 EJECTION SEAT ENERGY ABSORPTION SYSTEM

FOR REDUCTION OF CRASH FORCE LOADS, NADC Report ACEL-335,

Naval Air Development Center, Warminster, Pennsylvania, June 1957.

A; 119

* - - . . . . . . . . -... . * .
compressive loads; although this capab ity s still inadequate.

It is desirable that the tube elongate throughout its length

rather than locally; for example, at the end attachments. A

successful method of achieving nearle uniform elongation is the .

use of a low-modulus bonding agent between the tube and the appropriate end fitting (see Reference 66).

Typical load-elongation characteristics of a 0.02-in. wall by

0.50-in.-diameter stainless steel tube are illustrated in Figure 39.

e

Epoxy bonTest length

0.02-in. All x 0.50-in. diameter tube

3000 W 0,21 1ý eacl , .- in . --- _____

/< • Failure

S•.•_ 2000 - ... 'd

M 1500. Max dynamic Results based on

0 [/---Min dynamic 2 static and 12 f -Static I dynamfi(; tests

500 3044 ANNEALED STAINLESS STEEL ....

1 2 3 4 5 6 7

Deformation, in.

1 10 20 30 40 50 60

Deformation (elongation), percent

Figure 39. Comparison of dynamic and static load-elongation

curves for stainless steel tubes.

120

ME N_
5.7 ELONGATION OF BASIC STRANDED CABLE

This device has the same characteristics as the basic metal

tube or flat strap; however, the flexibility of a cable obviously has advantages for some load-limiter applications. The

cable end fittings are capable of sustaining the ultimate load

of the cable under static and dynamic conditions. This device

appears to be most applicable to bracing lightweight seats,

such as troop and gunner seats.

5.8 ROD PULL-THROUGH TUBE

This device uses the force required to expand the diameter of

a tube as a hardened, oversized rod or tube is drawn through

it, or to compress an inner rod or tube as it is drawn through

a die. The force required to overcome friction also contributes to the ernergy absorbed by this device; however, dependence

on friction to maintain a uniform load is unpredictable. The

frictional resistance of the device tested in Reference 66 (a

compression tube device with a rigid outer cylinder) was reduced by lubrication, but the device exhibited an initial peak

load as indicated by point A in Figure 40.

It can be seen in Figure 40 that the stroke of this device was

limited to 4 in. and that the failure load was about three

times the stroking (sustained) load. Thus, the tested device

had a safety factor of at least 3 to 1 built into it, and this

fact partially accounted for the poor specific energy rating

shown in Table 3. It can be seen in the figure that the maximum variation in the stroking load was from 1300 to 1600 lb,

or about 21 percent.

5.9 TUBE FLARING

This device simultaneously uses the forces required to expand

the diameter of a tube to the failure point and to bend the

tube walls through 90 degrees. The tube wall either shatters

into fragments or rolls up into spirals around the periphery

of the tube, as illustrated in Figure 41. A review of Reference 69 indicates that the above processes are sefisitive to

the ratio of the wall thickness to the die -radius; and that

ratios of less than 0.3 are likely to result in a rolling process, while ratios of greater than 0.4 are likely to result in

the fragmentation on the basis of tests using 2024-T3 aluminum

tubes.

This concept has been evaluated for an experimental crewseat

by the U. S. Naval Aircraft and Crew Systems Technology Directorate, now part of tbI' U. S. Naval Air Development Center, as

described in Reference 70. The device was used as the vertical

1 121

- ~~ - -- ---- - -*- - ---- - - - -- - - - - -
i.n

* .A

5000- 50Compression - tube .

Wt 0.86 lb each Failure

4500 CODE

Max dynamic Results based on

4000 - . Min dynamic 1 static and 4_ _ -0Static dynamic tests

3500

S3000

2500 Load discontinued - before bottoming

A occurred,

2000 ---- 4

I' I

1 5 0 0- - - - - - - " . .

- I! I .

I.000 " •____ _ _ __ _ __ _ _ _ __ _ _ _ I I

! .I

500ii I-- -___

0 3 .

0 2 3 4 5 6

Deformation (elongation), in.

Figure 40. Comparison of dynamic and static load-deformation

curves for compression tubes.

energy absorber in the seat. The device also was used as the

vertical load limiter for an experimental troop seat, as described in Reference 63.

The device cannot sustain rebound forces because only a minimum rebound resistance is provided by friction between the tube

and the forming die. However, a mechanism was installed in the

forming die to grip the tube against rebound movement.

122'

0
Forming die Fragmentation Rolling

Figure 41. Illustration of fragmentation and rolling- processes in tube-flaring device.

5.10 TENSION PULLEY

The tension-pulley that the pulley 

load limiter shown in Figure 42 requires

rotates. The casing be literally split open as the pulley

fects plastic deformation of the casing material efis unidirectional 

the energy absorption. 

and As the name implies, the device

has been used operates under tensile loading only. It in cargo restraint systems and energy-absorbing troop seats, as described in Reference 75.

5.11 SELECTION OF AN OPTIMUM LOAD LIMITER

An optimum load-limiting of the data presented system cannot be selected on the basis

lines with due 

above. The data should be used as guide- consideration to the requirements for each spa- cific application.

75. Turnbow, 

NAMIC TEST 

J. 

OF 

W., Robertson, S. H., and Carroll, D. F., DY- AN EXPERIMENTAL TROOP SEAT INSTALLATION AN H-21 HELICOPTER, Aviation IN

Division of Flight Crash Injury Research (AvCIR), Safety Foundation, nical Report Inc.1 TRECOM Tech- 63-62, U. S. Army Aviation Materiel Labora- tories, Fort Eustis, Virginia, November 1963.

123
'I

Deformed

housing

jihi AA

I I

Pigure 42. Tension-pulley load limiter.

124

• .. __ ... _.. .. i= , • ':
6. SEAT CUSHIONS

6.1 INTRODUCTION

The seat bottom and back with which the occupant is in constant

contact should be designed for comfort and durability. Sufficient clearance between fabric backs and bottoms or sufficient

cushion thickness of the appropriate material stiffness should

be provided to preclude body contact with the seat structure

when subjected to either the specified operational or crash

loads. Seat bottoms made of fabric should be provided with

I: means of tightening to compensate for sagging in use. The conflicting requirements of long-term comfort-versus-crash safety

considerations have made this a difficult design area.

For seat cushions, the problem is one of developing a compromise design that will provide both acceptable comfort and

safety. In the past, the comfort requirement was met by pro--

viding very thick, soft, foam cushions that allowed the occupant to sink in deeply, thereby producing a contour and

spreading the load around the person's buttocks so as to decrease local high pressure and eliminate point loading. This

approach provided both immediate and long-term comfort. 

A

method of providing thermal comfort was to force air through

4 the cushion, or to use stretched net cushions, which provided contouring and load spreading as well as the free passage of

air. The passage of air allows the evaporation of sweat and,

thus, achieves the desired cooling effect.

Crash-safety considerations require a minimal thickness of foam

to minimize or eliminate vertical motion of the pelvis during

high vertical loadings. This requirement conflicts with the

method chosen for providing pressure comfort described in the

previous paragraph, and constitutes a problem that must be

solved to provide an acceptable cushion.

One approach producing the desired compromise between crash

safety- and comfort uses a cushion hase-with a contour that

matches the universal buttocks configuration as closely as possible. This wraparound configuration spreads the load and decreases localized pressure without resorting to soft foams.

Additional comfort layers of foam can then be added to the base,

and the cushion base can be equipped with slots or holes which

allow for fore-and-aft passage of air to provide the desired

cooling. A layer of rate-sensitive foam can be used on top of

the base to provide a contour transition softer than the base.

This layer must either be open celled, or holes must be provided to allow for vertical movement of air. A layer of soft,

open-celled foam can be used on top of the rate-sensitive foam

to provide the initial comfort material and to also provide

125

-

1
vertical and horizontal air motion. The entire cushion can be

covered with a fire-retardant, open, nylon material to provide

for wear and abrasion resistance.

The total thickness of the compressed cushion at the buttock

reference point should be minimized and can be limited to between 0.5 and 0.75 in. of thickness. This cushion probably permits not more than 3/8 in. of vertical motion of the pelvic

structure (ischial tuberosities) from the 1-G loaded position

to the full vertical crash-loaded position.

Other methods of achieving the desired effect are available.

One is to include the basic provisions just described but to

achieve the thermal effect plus some loading comfort by the

use of special coverings such as lamb's wool. This type of

cover uses the lamb's skin with a small depth of combed and

clipped wool on the occupant interface surface. These covers

need holes cut through the leather to allow free passage of

air for cooling as previously discussed.

To meet the required crashworthy characteristics, the optimum

aircraft seat cushion will:

"* Be extremely lightweight.

"* Possess flotation capabilities.

"* Be nonflammable.

"• Be nontoxicl will not give off fumes when burned,

charred, or melted.

"* Be tough and wear resistant.

"* Be easily changeable.

"* Provide comfort by distributing the load and reducing

or eliminating load concentrations.

"* Provide thermal comfort through ventilation.

"* Provide little or no rebound under crash loading.

"* Allow an absolute minimum of motion during crash

loading.

6.2 REQUIREMENTS

For seats of light movable weight (less than 30 lb), cushions

should be used for comfort only. The maximum uncompressed

126
thickness for a properly contoured cushion should be 1-1/2 in.,

unless it can be shown through analysis or through dynamic

tests that the cushion design and material properties produce

a beneficial (reduced force transmissibility) result.

For seats of greater movable weight, such as integrally armored seats, every effort should be made to design a cushion

that minimizes relative motion between the occupant and the

seat and that acts as a shock damper between the occupant and

* the heavy seat mass. Viscoelastic and loading-rate-sensitive

materials, such as discussed previously, can be used to accom-

* plish this goal. Again, dynlamic analysis and/or testing should

be conducted to demonstrate that the cushion design produces a

desirable system result over the operational and crash environmental range of interest.

6.3 ENERGY-ABSORBING CUSHIONS

The use of load-limiting cushions in lieu of load-limiting

seats is undesirable for two reasons:

a The downward movement of the torso into a crushable

seat cushion produces slack in the restraint harness.

This slack could allow injury during subsequent longitudinal or lateral acceleration in forward-facing

seats by contributing to dynamic overshoot and/or by

allowing the lap belt to move upward into the soft

portion of the abdomen. For an aft-facing seat, this

slack is not as significant for longitudinal accelerations but applies to the lateral direction.

* A crushable cushion does not make optimum use of the

available stroke distance since space must be allowed

for the crushed material. A crushable cushion can

be only approximately 75 percent as efficient as a

mechanical load-limited system thAt allows the seat

to stroke completely to the floor.

Crushable cushions are impractical in rotary- and light fixedwing aircraft because of the long stroke distance required

to attenuate the high vertical - loads present in the 95thpercentile crashes. The only justifiable use of energyabsorbing cushions instead of load-limited seats might be in

retrofit circumstances where, because of limitations in existing aircraft, another alternative does not exist (see Reference 76 for further information on energy-absorbing cushions).

76. DYNAMIC TEST OF CRUSHABLE SEAT CUSHIONS, AvSER Report

M67-6, Aviation Safety Engineering and Research (AvSER),

Division of Flight Safety Foundation, Inc., Phoenix, Arizona, August 1967.

127

low• -o. .. .,.N. ,..
.. 1ý7 "Wr 

Recent research has indicated that foams can be used more eco

nomically than honeycombs without reduction in performance.

Foams are much easier to form and are less costly than metallic

honeycomb materials and are therefore recommended for this use.

6.4 NET-TYPE CUSHIONS

This type of cushion serves the same purpose as the filled

cushion; however, a net material is stretched over a contoured

seat frame, and the body is supported by diaphragm action in

the net rather than by deformation of a compressible material.

The net-type cushion might more properly be called a net su -

port. If a net support is used in the seat, its reboud caacteristics should be capable of limiting the return movement

from the point of maximum deformation to 1-1/2 in. Net supports should not increase the probability of occupant submarining or dynamic overshoot. The net elastic-stretch limitation might be achieved by including a stiffer net, such as a

steel or aluminum woven material under the net support.

6.5 OTHER CUSHIONS

In most cases the back cushion will not play a significant role

in the crash dynamics; however, it will influence comfort and

can influence the injury tolerance of the spine. The cushior

should be of a lightweight foam material or net. The foam can

be a standard furniture type that meets the other requirements

listed in Section 6.2. Lumbar supports, particularly those

that are adjustable by the occupant, are desirable for comfort

and for safety reasons. A firm lumbar support that holds the

lumbar spine forward in extension increases the tolerance to

+G loading. z

6.6 HEADRESTS

A headrest should be provided for occupant head/neck whiplash

protection. Headrest cushions are used only to cushion head

impact and prevent whiplash injury due to backward flexure of

the neck. The cushioning effect can be provided by a thin pad

and a deformable headrest or a thicker cushion on a more rigid

headrest. For the thicker cushion, the provisions of Section

10.9 should be applied and at least 1.5 in. of cushion should

be provided if possible within the space limitations of the

application.

128

.... .
I

7. DESIGN PRINCIPLES FOR PERSONNEL RESTRAINT SYSTEMS

7.1 INTRODUCTION

Crash injury accident statistics indicate that failure of personnel restraint harnesses has been a frequent cause of injuries and fatalities in U. S. Army aircraft accidents. This is

unfortunate because body restraint is relatively easy to control. Adequate restraint in a crash can mean the difference

between life and death, since evacuation from a burning or

sinking aircraft is considerably improved if no prior injury

or debilitation has occurred. It is the intent of this section to provide general criteria and quidelines for the design

of personnel restraint systems to reduce injury or debilitation

in a crash situation. Design criteria for cargo restraint systems are presented in Volume III.

Restraint harnesses for personnel should provide the restraint

necessary to prevent injuries to all aircraft occupants in

crash conditions approaching the upper limits of survivability.

Appropriate strength analysis and tests as described in Section

8.5 should be conducted to ensure that a restraint system is

acceptable.

*1 Numerous methods of restraining the human body have been proposed, investigate&, and used. Some of these have proven to

be exceptionally good, and some have left much to be desired.

However, there are certain qualities that a harness should possess if it is to be used routinely for military flights. These

desirable qualities are listed below:

* Comfortable and light in weight.

* Easy for the occupant to put on and take off even in - the dark.

* Contain a single-point release system, easy to operate with one (either) hand since a debilitated person might have difficulty in releasing more than one

buckle with a specific hand. Also, it should be protected from inadvertent release, e.g., caused by the

buckle being struck by the cyclic control or by inertial loading.

e Provide personnel with freedom of movement to operate

the aircraft controls. This requirement necessitates

the use of an inertia reel in conjunction with the

shoulder harness.

129
1 I

# Provide sufficient restraint in all directions to

prevent injury due to decelerative forces in a po- tentially survivable crash.

* Webbing should provide a maximum area, consistent

with weight and comfort, for force distribution in

the upper torso and pelvic regions and should be of

low elongation under load to minimize dynamic overshoot.

7.2 TYPES OF SYSTEMS

7.2.1 Aircrew Systems

The existing military lap belt and shoulder harness configuration with a center tiedown strap as shown in Figure 43 is the

minimum acceptable harness for use by U. S. Army pilots. The

lap belt tiedown strap resists the upward pull of the shoulder

straps and prevents the belt's displacement into abdominal tissue. The tiedown strap is comfortable to wear since it does

not contact the pelvis, and it is narrow enough within limits

of acceptable strength so that little leg rubbing is encountered by the wearer during antitorque or rudder pedal operation.

The configuration shown in Figure 44 provides improved lateral

restraint due to the addition of the reflected shoulder straps. This system, which resulted from the investigation reported in

Reference 77, consists of one dual-spool inertia reel or two

separate inertia reels with two reflected straps, a shoulder

harness collar assembly, a lap belt assembly including retractors, and a buckle assembly. The buckle assembly consists of

a single-point release buckle permanently attached to the tiedown strap. The tiedown strap consists of a fixed-length strap

for any specific seat and cushion design, and an anchor fitting

that connects the strap to the seat pan beneath the seat cushion. The left- and right-hand lap belts, connected at the

single-point release buckle, are attached to the seat or-aircraft structure through automatic rock/unlock retractors.

The shoulder harness collar assembly consists of a pad in the

form of a collar fitting around the crewman's neck, over which

the shoulder harness straps are routed. The lower shoulder

77. Carr, R. W., and Desjardins, S. P., AIRCREW RESTRAINT SYSTEM - DESIGN CRITERIA EVALUATION, Dynamic Science, Division of Ultrasystems, Inc.1 USAAMRDL Technical Report 75-2,

Eustis Directorate, U. S. Army Air Mobility Research and

Development Laboratory, Fort Eustis, Virginia, February

1975, AD A009059.

130

- -!
I

Item identity.3

1. Buckle assembly

A. Single-point

release buckle 3B

B. Tiedown strap C. Tiedown anchor

2. Lap belt assembly

A. Lap belt

Adjuster 

3. Shoulder harness

assembly

A. Inertia reel

B. Inertia reel strap

C. Lower shoulder

D. Adjuster strap 3C!

2I

1B A 4

ardI

Figure 43. Basic aircrew restraint system.

•131

" "A• I'i

~rd

m m • •• " • " • • "I
item identity•

1. Suckle assembly

A. Single-point

B. Tiedown strap

C. Tiedown anchor B

2. Lap belt assembly

A. Lap belt

B. Retractor

3. Shoulder harness

collar assembly 3

A. Pad

B. Roller fitting C. Adjuster

D. Lower shoulder strap i"

4. Inertia reel assembly

A. Reflected strap

B. Anchor D

I "

~rd

Figure 44. Aircrew restraint system, including

reflected shoulder straps.

132

S1 !OWN
straps connect to the bottom of the collar assembly through the

adjusters. The reflected straps pass through the roller fittings at the top of the collar. Each reflected strap is extended forward from an inertia reel, looped through the roller

fitting, and then directed rearward to the opposite side of the

seat back. These straps are attached to the seat through anchor fittings on the reflected ends and through inertia reels

at the other end. The lap belt straps, tiedown strap, and

lower shoulder straps are all connected at the single-point

release buckle. Details of the hardware in these systems are

discussed in Section 7.5.

7.2.2 Troop Systems

Considerations in the selection of a troop or passenger seat

restraint system are different from those for an aircrew system. First of all, the seat may face forward, sideward, or

aftward. Secondly, the restraint system must be capable of

being attached and removed quickly in an operational environment by troops encumbered by varying types and quantities of

equipment. Also, whereas a pilot probably u..es the restraint

system in his aircraft so frequently that its use becomes a

matter of habit, troops and passengers can be expected to be

unfamiliar with the system. The effects of this lack of familiarity would probably become more pronounced in a combat situation when the risk involved in not using the restraint system

becomes even higher. Therefore, hardware should be uncomplicated and, if possible, resemble the familiar, such as automotive hardware. Finally, the need to quickly remove and stow

the seats requires compact and lightweight restraint systems.

For the aft-facing passenger, the need for a tiedown strap is

negligible, since the seat back will provide the primary restraint; however, the shoulder harness should be retained as

part of the restraint system to provide adequate support in

crashes that produce significant lateral loads.

It is difficult to provide adequate restraint for stde-facing

passengers with a lap belt and shoulder harness alone. Leg

restraint would also be preferred; -however, leg restraints are

generally not practical because of operational requirements. which necessitate the side-facing seats in Army aircraft. A

reflected shoulder strap and side belt strap offers a compromise solution; however, they too have met with resistance because of weight and cost considerations. The belt side straps,

extending from the lap belt high on the thigh to the seat pan

forward of the lap belt anchor, help to hold the belt in place

over the pelvic region as well as provide more area to resist

the pressure from the pelvis. The reflected shoulder strap

provides improved upper torso restraint.

133

wI
Two systems that resulted from the investigation reported in

Reference 78 are shown in Figure 45. The Type II troop restraint system was designed to mount on a forward-facing or

aft-facing troop seat and consists of a two-strap shoulder harness and a lap belt assembly. The two shoulder straps are attached to two single inertia reels. They extend forward and

down over the occupant's upper torso and are connected into

the single-point release, lift-lever buckle. The lap belt assembly includes left- and right-hand belts, with adjusters,

that are connected together at the lap belt buckle. The Type

I troop restraint system was designed to mount on a side-facing

troop seat and differs from the Type II restraint by having a

single shoulder strap that passes diagonally across the occupant's upper torso. It should pass over the shoulder closest

to the nose of the aircraft. If the Type I system is used in

either a forward- or aft-facing seat, the diagonal shoulder

strap should pass over the outboard shoulder to restrain the

occupant from protruding outside the aircraft during lateral

loading.

7.2.3 Crew Chief and Door/Window Gunner •ystems

Restraint systems for crew chiefs and door/window gunners are

similar to troop systems; however, they must allow the crewmember to move out of the seat to perform duties such as maneuvering the gun or observing tail rotor clearance while landing

in unprepared areas. The system should restrain the occupant

to the seat the instant he returns to the seat and provide adequate restraint during a crash. The system should maintain

the lap belt buckle in the proper relationship to the gunner,

preventing the shoulder straps from pulling it up or the lap

belt from pulling it sideways. Such a system has been described in Reference 79 and is shown in Figure 46. It consists

of a lap belt with inertia reels on each side of the seat and

two shoulder straps connected in an inverted-Y arrangment to a

single inertia reel strap. The lap belt with thigh strap attachment is easy to put on and prevents the lap belt from riding up during operation of the gun. The lap belt is plugged

into the two seat pan inertia reels when the crewmember is:to

78. Carr, R. W., HELICOPTER TROOP/PASSENGER RESTRAINT SYSTEMS

DESIGN CRITERIA EVALUATION, Dynamic Science, Division of

Ultrasystems, Inc.; USAAMRDL Technical Report 75-10, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, June 1975,

AD A012270.

79. Reilly, M. J., CRASHWORTHY HELICOPTER GUNNER'S SEAT INVESTIGATION, The Boeing Vertol Company; USAAMRDL Technical

Report 74-98, Eustis Directorate, U. S. Army Air Mobility

Research and Development Laboratory, Fort Eustis, Virginia, January 1975, AD A005563.

134
22

22

SI ',

2, / 4.._ /s

I" [ . I\- -3

66

Type I Type II

Item identity

1. Inertia reel

2. Shouldar strap

3. Lap belt anchor

4. Buckle with shoulder strap connection

5. Lap belt

6. Adjuster/fitting

Figure 45. Aircraft troop/passenger restraint systems.

135

;t, .- - --. - .:- .-. A , , . .-
. II

5. La.bl

-i

4

Itenm identity

2. Shoulder strap

3. Shoulder strap adjuster 4. Attachment release buckle

6. Lap belt inertia reel

7. Thigh straps * I

8. Thigh strap adjuster 9. Lap belt plug-in fitting

\

Figure 46. Gunner restraint system. (Prom Reference 79)

136

______.....__ __..... ..... I
be seated or standing in front of the seat. The shoulder harness and lap belt with thigh straps may serve as a "monkey harness" when the crewmember disconnects the two lap belt plug-in fittings from the inertia reels. The resultant configuration

permits the crewmember more extensive travel within the cabin

while still being connected to the shoulder harness inertia

reel, thereby restraining the crewmember from falling out of

the aircraft.

7.2.4 Inflatable Systems

An automatically inflatable body and head restraint system for

helicopter crewmen has been jointly developed and tested by

-. .the Naval Air Development Center and the Applied Technology

"Laboratory. As illustrated in, Figure 47, this system provides

increased crash protection because it provides automatic pretensioning that forces the occupant back in his seat, thereby

reducing dynamic overshoot and reducing strap loading on the

wearer when the inflated restraint is compressed during-the

crash. The concentration of strap loads on the body are reduced because of the increased bearing surface provided when

the restraint is inflated, and both head rotation and the possibility of whiplash-induced trauma are also thus reduced.

Although more complex and costly than conventional belt systems, such a system may be justified because of its occupant protection potential. Development of the system and results

of testing are documented in References 80 and 81.

7.3 GENERAL DESIGN CRITERIA

7.3.1 Comfort

Comfort must not be compromised by crash-survival requirements

for obvious reasons. For example, a lap belt with an adjustment fitting located directly over the iliac crest bone would

provide a constant source of irritation that would result in

eventual fatigue to the wearer. The main comfort consideration for restraint harnesses is-the absence of rigid hardware

80. Schulman, M., and McElhenney, J., INFLATABLE BODY AND HEAD

RESTRAINT, NADC-77176-40, Naval Air Systems Command, Peartment of the Navy, Washington, D. C., September 1977, D A046477.

81. Singley, G. T., III, TEST AND EVALUATION OF IMPROVED AIRCRAFT RESTRAINT SYSTEMS FOR COMBAT HELICOPTERS, Paper No.

A.18, presented at NATO/AGARD Aerospace Medical Panel,

Aerospace Specialist's Meeting on Aircrew and Survivability, North Atlantic Treaty Organization, Bodo, Norway,

May 20-23, 1980.

137-.
I -4

II

1381

0i

I

i*
located over bony portions of the torso. Also, webbing that

is too wide or too stiff could cause discomfort.

7.3.2 Emergency Release Requirements

From a crash survival point of view, it is mandatory that a

shoulder harness/lap belt combination have a single point of

release that can be operated by one (either) hand to make it

easier for debilitated occupants to quickly free themselves

from their harnessing in a severe crash because of the dangers of postcrash fire or sinking in water. The force required to

release the harness with only one finger should fall between

20 and 30 lb on the basis of existing requirements for military harnesses. An excessive force could hii.der rapid emergency release, while a light force could cause inadvertent release. Further, the release should be possible with the weight

of the occupant hanging in the restraint system after experiencing the full crash loads. This will guarantee that the

occupant can release the system after a severe survivable crash

even when inverted in a loaded restraint system. The release

forces for the inverted case should be minimized and, in any

case, should not exceed 50 lb applied with only one finger. It

should be possible to produce the torque necessary to release

rotary buckles by applying a load at a single point on the

handle as described above.

In restraint systems other than the Type I of Figure 45, if a

lift latch or similar type buckle is used, the restraint system

design should ensure that the latch lifts from left to right on

all installations. This will reduce the possibility of reverse

installations and their resultant hazard.

The release device must either have the capability to withstand the bending moments associated with doflections and motions during loading, or it should contain features that allow

the fittings to align themselves with the loads, thereby reducing or eliminating the moments. If belt loading direction is

such as to cause the strip to-bunch up in the end of a slot,

failure can occur through initiation of edge tear. As a result of an investigation of restraint system design criteria

reported in Reference 77, the fitting angles illustrated in

Figure 48 are recommended.

Eliminating fitting rotation in the flat plane of the buckle

during loading may prove to be difficult in lightweight systems. If the integrity of the attachment of the fitting within

the buckle can be compromized by rotation, then rotation must

be completely eliminated. Experience has shown that it is

better to design the attachment of the fitting within the

139
-it

9

:4

Figure 48. Buckle fitting attachment and motion angles.

buckle to be insensitive to rotation, i.e.,r a round pin in a

round hole, rather than a flat-faced dog which must seat on a

flat face of a slot, than to rely on restraining the fitting

against rotation. In the latter case, a small amount of rotation can cause point loading of a corner of the dog against

one end of the slot. The point loading can easily increase

the stress applied at the contact point to i..s ultimate bearing strength. This will result in metal deformation and the

formation of a sloped surface which then can act to cam open

the attachment mechanism.

Further, the release mechanism (buckle) should be protected

against accidental opening. Neither decelerative loading of

components nor contact with aircraft controls such as cyclic

controls should open the device. It was mentioned earlier in

this volume that required cockpit dimensions should be reviewed. It appears that the occupant can be placed too close

to the cyclic control in helicopters and that a fully retracted

cyclic head can contact the buckle. The buckle release mechanism should be protected against inadvertent release either

during operation or in-a crash. It should be emphasized that,

iof contact between the cyclic control and the buckle is possible in an-operational mode, a considerable overlap can exist

during crash loading when the restraint system is deformed forward several inches.

7.3.3 Lap Belt Anchorage

The anchorage points for the lap belt can be locatei either on

the seat bucket or on the basic aircraft structure. if the anchorage is located on basic aircraft structure, the movement

of the seat under the action of load-limiting devices must be

140

I ,, , , -. . '" . .. • .. " ----- - •-, ,•• : • f • III i•' --- d
It

considered to ensure that the lap belt restraint remains effective regardless of seat position. If the seat includes longitudinal load limiting, attachment of the lap belt to the basic

structure will not be practical because then most of the for-.

ward load will be car.ied to the aircraft structure through the

belt rather than through the seat. In such a case careful consideration must be given to the belt strength since the belt

must restrain the motion of the seat, as well as the occupant.

The lap belt should be anchored to provide optimum restraint

for the lower torso when subjected to eyeballs-out (-G )

forces. One of the anchorage variables which has an infl~ence

on restraint optimization is the location of the lap belt anchorage teristic in is the the fore-and-aft angle in a vertical direction. fore-and-aft The important plane charac- between I

a projection of the lap belt centerline and the buttock reference line, or plane. This angle defines the geometrical relationship between the longitudinal and vertical components of

the belt load. A small angle provides an efficient path for

supporting longitudinal loads while a large angle provides an

efficient system for supporting large vertical loads. Thus,

for supporting large forward-directed loads, a small angle

would be desirable, but for reacting the large vertical loads

imposed on the lap belt by the loaded shoulder harness a large

angle is required. The compromise for location of the anchorage must consider all the variables including the tendency

for the occupant to submarine under the lap belt. In an accident with high combined vertical and longitudinal impact forces

the restrained body will tend to sink down into the seat (where

the magnitude of the displacement depends on cushion properties) and almost simultaneously be forced forward. This movement is illustrated in Figure 49. If the lap belt angle is too

small the belt can tend to slip over the iliac crests of the

pelvic bone, allowing the pelvis to rotate under the belt. The

inertial load of the hips and thighs tend to pull, or submarine, the lower torso under the belt. Lower torso restraint

is then accomplished through lap belt loading of the soft abdominal portions of the body, possibly causing visceral injury

in addition to the spinal injury illustrated in Figure 49.

To counteract the tendency for submarining the lap belt angle

can be increased; however, the load in the belt increases for

a given torso deceleration because of the smaller longitudinal

loading component available to restrain the occupant. Further,

additional forward motion is allowed because of the increased

deflection of the webbing caused by the increased loading and

the greater forward rotation of the lap belt. However, as the

webbing is loaded, it presses down into the thighs of the occupant and minimizes the possibility of submarining by picking

141

--
P' Sioulder !,arno~'s r'all

----- --.. . ... . . .. . .. .- • ' -

* d ~ lxn.ctuo1 torso dis'laccfrnt

Figure 49. Pelvic rotation and submarining caused by

high longitudinal forces combined with

moderate vertical forces.

up the longitudinal component of restraint load required to

bring the system into equilibrium.

In order to avoid the increased possibility of both spinal and

abdominal injury, a properly designed restraint system should

not allow submarining to occur. Still, an efficient angle

should be maintained to limit the forward motion of the occupant.

Comfort is another concern in lap belt anchor location. A pilot must raise and lower his thighs during operation of rudder

pedals or antitorque pedals. If the lap belt anchor is too far

forward, the lap belt will pass over the pilot's thighs forward

of the crease between the thighs and the pelvis and thus may

interfere with vertical leg motion. It is important, therefore, to position the lap belt anchorage so that it provides -

optimum restraint while not interfering-with the pilot's operational tasks. A forward location of the anchor does not negatively influence the comfort of passengers since passengers

are not required to perform operations with their legs.

In order to accomplish these objectives, the vertical angle

between the lap belt centerline and the buttock reference line

as installed on the 50th-percentile occupant should not be less

than 45 degrees and should not exce 5 degrees, as shown in

Figure 50(a). Further. it is deF Ae to locate the anchor

point at or below the buttock r, Bnce line for comfort and

142
a) All systems b) Systems with lap belt c) Systems without tiedown tiedown straps straps

4 5 -5 5V4

Seat reference L

uttock point ;typical) "7

reference

line 2.0 + .25 - 0

Figure 50. Lap belt anchorage geometry.

performance. If the anchor point must be located above the

buttock reference line, as on most armored seats, the anchor

point should be positioned to ensure that the belt-angle lies

within the desired 45- to 55-degree range. For a system having

a lap belt tiedown strap to counteract the upward force of the

shoulder harness (e.g., in pilot seats), the lap belt anchors

should be positioned so that the centerline of the lap belt

passes through the seat reference point as shown in Figure

50(b). If the restraint system does not have a tiedown strap

(e.g., in passenger seats), the lap belt anchor should be positioned so that the belt centerline passes through the buttock

reference line 2 to 2-1/4 in. forward of the seat reference

point as shown in Figure 50(c). This position provides sufficient vertical load components to help counteract the upward

force of the shoulder straps. For positioning anchors that do

not fall on the buttock reference line, the angle between the

lap belt centerline and the buttock reference line can be assumed to be 45 degrees for systems with tiedown straps and 55

degrees for those without.

Submarining can be reduced by ensuring that the lap belt is

tight, as shown in studies reported in Reference 82. Thus,

82. Roberts, V. L., and Robbins, D. H., MULTIDIMENSIONAL

MATHEMATICAL MODELING OF OCCUPANT DYNAMICS UNDER CRASH

CONDITIONS, Paper No. 690248, Society of Automotive Engineers, Inc., New York, January 1969.

143

-. -,
!I

* 

-'

-,1

care should be taken to train occupants to tighten the lap belt

to the maximum consistent with comfort and to not loosen the

belt anytime during flight.

For seats that limit lateral motion of the occupant with structure, such as in armored seats, the anchorage point and hardware should possess sufficient flexibility and strength to sustain design belt loads when the belt is deflected laterally

toward the center of the seat through an angle of up to 60 degrees from a vcrtical position. The side motion of fittings on

other seats should also be capable of supporting design loads

with the lap belt deflected laterally away from the center of

the seat through an angle up to 45 degrees from the vertical.

These recommendations are made to ensure that lateral loading

on the torso will not result in lap belt anchorage failure.

7.3.4 Shoulder Harness Anchorage

The shoulder harness or inertia reel anchorage can be located

either on the seat back structure or on the basic aircraft

structure. In placing the inertia reel, strap routing and pos- sible reel interference with structure during seat adjustment

or energy-absorbing stroke of the seat must be considered. Location of the anchorage on the basic aircraft structure will

relieve a large portion of the overturning moment applied to

the seat in longitudinal loading; however, due consideration

must be given to the effect of seat bucket movement in loadlimited seats. Vertical movement of the seat pan can be pro- vided for by placing the inertia reel aft of the seat back

shoulder strap guide a sufficient distance so that seat vertical movement will change the horizontal position and the angle

of the straps very little.

Shoulder straps should pass over the shoulders in a plane per- pendicular to the back tangent line or at any upward (from

shoulders to pull-off point) angle not to exceed 30-degrees,

as illustrated in the upper-left sketch in Figure 51.

Any installation that causes the straps to pass over the shoulders at an angle below the horizontal adds additional compressive force to the seat occupant's spine as shown in the lower

sketch of Figure 51. A shoulder harness pull-off point at

least 26 in. above the buttock reference line is needed to ensure that the straps do not apply an excessive downward load on

the spine of a 95th-percentile occupant.

The shoulder harness anchorage or guide at the top of the seat

back should permit no more than 0.5-in. lateral movement (slot

no more than 0.5 in. wider than strap) to ensure that the seat

occupant is properly restrained laterally. The guide should

provide smooth transitions to the slot. The transition contour

144

4.
Correct

Backtanent ineTorso carries small

M~aximum vertical load

shoulder

strap

angle, -. Shoulder

strap load

v, (Force diagram)

2 6.0 to

271.0 in.k

Seat referec

point

Buttock reference / b~ s

line -

0)

Incorrect i

Torso carriei-large vertical load

Figure 51. Shoulder harness anchorage geometry.

should be of a radius no less than 0. 25 in. and should extend

completely around the periphery of the slot to minimize edge

wear on the strap and reduce the possibility of webbing failure

due to contact with sharp edges under high loading.

7.3.5 La etTeon ta nhrg

When the upper body is thrown forward against the shoulder

straps, an upward pull is exe' - on the lap belt. If a lap

..... ........ 
I

belt tiedown strap is not being used, the tendency is for the

belt to be pulled up off of the iliac crests and into the soft

solar plexus area, causing injury to abdominal viscera as pre- viously shown in Figure 49. A tiedown strap attached to the

center of the lap belt will prevent the upward belt movement.

It is recommended that the tiedown strap be located on the seat

pan centerline at a point 14 to 15 in. forward of the seat back. For shorter seat pans, the anchor must be placed as far forward

as possible.

7.3.6 Adjustment Hardware

Adjusters must carry the full design load of the restraint sys- i

tem subassembly, of which they are a part, without slipping or

crushing the webbing between items such as locking cams and

the opposite locking surface. In extremely highly loaded applications, this may require that the strap be doubled in a

manner that requires the adjuster to carry only half of the

strap assembly load. The force required to adjust the length

of webbing should not exceed 30 lb in accordance with existing

military requirements for harnesses. Insofar as possible, all 3

adjustments should be easily made with one (either) hand. Adjustment motions should be toward the single-point release

buckle.

7.3.7 Location of Adjustment and Release Hardware

This hardware must not be located directly over hard points of

the skeletal structure, such as the iliac crests of the-pelvis

or the collarbones. The lap belt length adjusters should be

located either at the center of the belt near the release

buckle or at the side of the hips below the iliac crests, preferably the latter. The shoulder strap adjusters should be

located as low on the chest as possible in order to avoid concentrated pressure on the collarbones.

7.3.8 Webbing Width and Thickness Requirements

Selection of the optimum webbing width for a lap belt and

shoulder harness must be based on two conflicting requirements:

(1) maximum width for lowest pressure and (2) minimum width for

maximum comfort and minimum hardware weight. Webbing requirements are discussed in detail in Section 7.4.

7.3.9 Hardware Materials

All materials used for the attachment of webbing (release

buckles, anchorages, and length adjusters) should be ductile

enough to deform locally, particularly at stress concentration

points. Ductility in restraint harness hardware is not as critical when energy-absorbing provisions are incorporated into

146

- -- ....- .- -... . .. .......-- .. ..-- •- -
the seat, because the maximum loading of the system is limited.

Thus, it would be possible to specify low-ductility materials

on load-limited seats and to specify high-ductility, moderatestrength materials on nonload-limited seats. Such a specification could possibly lead to the inadvertent installation of

low-ductility harness fittings on rigid, nonload-limited seats.

For example, it is known that 20-G-strength shoulder straps have been mistakenly installed in place of 40-G straps. To

prevent such a possibility, it is concluded that, wherever applicable, all harness fittings should be made of equivalent high-ductility materials to ensure their interchangeability. A minimum elongation value of 10 percent (as determined by

standard tensile test specimens) is recommended for all metal

Sharness-fitting materials. The 10-percent elongation value

can be achieved with copper-base aluminum alloys, low-carbon

steels, and stainless steel. There are obviously some components that, for operational purposes, rely on hardness. These

components should be designed to perform their necessary function but be made from materials as nearly immune as possible

to brittle failures.

7.3.10 Structural Connections

7.3.10.1 Bolted Connections: Safety margins of 15 and 25 percent for shear and tensile bolts, respectively, are recommended

by most aircraft companies for the manufacture of basic aircraft structure. These margins are intended to allow for misalignment of holes, stress concentrations, and fatigue strength

reductions; however, the bolt's fatigue strength is not a factor for a one-time maximum loading as occurs in a crash. Thus,

it is concluded that the safety margins for shear and tensile

bolts in restraint systems can be reduced to 5 and 10 percent,

respectively.

Good aircraft engineering practice also dictates that bolts

less than 0.25 in. in diameter should not be used in tensile

applications because of the ease with which these smaller bolts

can be overtorqued. Wherever possible the bolts should be designed for shear rather than tension.

7.3.10.2 Riveted Connections: Riveted joint design guidelines

are presented in MIL-HDBK-5, "Metallic Materials-and Elements

for Aerospace Vehicle Structures" (Reference 23). This handbook i s recommended as a guide for restraint system hardware -

design.

7.3.10.3 Welded Connections: Welded joints can be 100 percent efficient; however, they may be only 50 percent efficient,

depending upon the skill of the welder. Since welded joints

can be completely acceptable and in some cases superior to

147

S k---4L .. . . : - .. - ., :•? *'-:_ 
bolted or riveted joints, it is not reasonable to prevent the

use of this type of joint if strict inspection procedures are

used to ensure that all welded joints are adequate. Welding

processes are discussed in Military Specifications MiL-W-8604,

-6873, -45205, and -8611. These specifications should be used

as guides to ensure quality welding.

Welded joints may contain stress concentration points and misaligned parts in a manner similar to bolted joints; therefore,

the cross-sectional area of the basic material in a welded

joint should be 10 percent greater than the area needed to sustain the design ultimate load.

7.3.10.4 Plastic Strength Analysis: Plastic analysis methods

should be used for strength determination wherever applicable

in order to obtain maximum-strength hardware at the lowest possible weight. Plastic analysis makes maximum use of the strain

energy available in ductile metals. References 28 and 29 cover

this subject.

7.4 WEBBING AND ATTACHMENTS

7.4.1 Properties

The maximum load to be sustained by restraint harnesses can be

determined by a review of seat load-deflection requirements

(Chapter 8). The curves shown there include the effects of

dynamic overshoot loads. The maximum load shown is 35 G for

the cockpit seat, where the seat structure provides for little

elongation. The required load is reduced as the deformation

is increased. Although the restraint harness could be designed

to varying loads in accordance with the energy-absorber G level

used in the seat, it is believed to be more practical and fool

proof to design a single-strength restraint harness that can be

interchanged with all seats of similar configuration and orientation. The main advantage of a single-strength harness would

be the assurance that it could be interchanged between loadlimited seats and nonload-limited seats without fear that an

understrength harness might be installed. On this premise,-the

design strength of all forward-facing and side-facing~restraint harnesses should be equal to or greater than the strength of

the cockpit seats because the higher design strength of the

cockpit seats would govern. At first, this solution might

seem to be too conservative because of the lower load levels

required for cabin seats; however, closer scrutiny indicates

that the assymmetrical nature of the forces on the harness in

the side-facing seats could result in loads just as high as

those experienced in the forward-facing cockpit harness for a

more symmetrical loading.

148

A __________11__________nln________I_____i-n
The distribution of the total load on the various harness components is not easily determined; however, these forces have

ueen fairly well approximated by theoretical calculations and

by experimental test data. The test data have been obtained

from tests on restrained 95th-percentile anthropomorphic dummies under a variety of test conditions. The maximum design loads for the various harness components attached on the seat 7

are listed in Table 4. These loads may appear to be higher

than necessary to offer restraint on a 35-G seat for a 222-lb

"occupant; however, these loads allow for (1) torso variations,

(2) asymmetric loadings, and (3) a safety factor to ensure that

the harness does not fail before the seat fails.

TABLE 4. RESTRAINT HARNESS COMPONENTS LOAD-ELONGATION

DESIGN AND TEST REQUIREMENTS (MIL-S-58095(AV))

Maximum elongation

Minimum load (design goal)

Harness components (lb)(a) (in.)(b) -

Inertia reel strap(s) 6 0 0 0 (c) (e)

Shoulder harness strap(s) 4 0 0 0 (d) 1 .5

Lap belt 4000 2.0

Lap belt tiedown strap 4500 0.5

NOTES: (a) Applied in straight tension.

(b) Total length of harness component tested must be

the same as when installed on the seat and adjusted for a 95th-percentile clothed occupant. (c) This represents the total load from all shoulder

straps. A single diagonal shoulder strap should

carry 6000 lb.

(d) This represents the minimum load that one of two

shoulder straps should carry. (e) This applies only to the shoulder harness and inertia reel strap outsiLde-the reel (exclusive of

the webbing wound on the spool of the inertia

reel).

It will be noticed that the inertia reel strap is required to

carry 6000 lb when the inertia reel itself is designed to carry

only 4000 lb. Since no inertia reel spool failures have occurred, and since combinations of shoulder strap loads exceeding 6000 lb have been measured, it must be assumed that the

loads in the straps are reduced thrnugh friction with the body

a and guide slots to no more than 4000 lb at the reel.

149
The elongation of all webbing used in the harness must be minimized to decrease overshoot. Dynamic tests conducted with anthropomorphic dummies and several tests with cadavers indicate

that a total elongation greater than about 1.5 in. will result

in an overshoot iad that is more than 25 percent above input

peak loads at tY.o floor. The load and deformation requirements indicated for st.is in Chapter 8 are based on the use of a restraint harness that elongates only 1.5 in. away from the seat

back. Torso compression and "normal" harness slack, however,

account for another 2 in. (at a 35-G load). Therefore, the

total torso movement away from the seat back is about 3.5 in. . I

The value of 1.5 in. for harness elongation is less than half

of the total torso movement. A lower value would not be reaaonable unless the normal harness slack and tissue deformation

are also reduced. Table 4 shows that the shoulder strap elongation is restricted to 1.5 in., while the lap belt is re- stricted to 2.0 in. of total end-to-end stretch or 1.0 in. of(

loop elongation. Restraint systems for the new generation of

Army helicopters use a low-elongation polyester webbing, the

characteristics of which are listed in Table 5.

TABLE 5. RESTRAINT WEBBING CHARACTERISTICS

Nominal Minimum

Restraint webbing Webbing brek.ing

system width thickness streigth Elongation*

component (in.) (in.) (lb) (percont)

Inertia

reel 1-3/4 0.057 6980 6.9 @ 3000 lb

Shoulder

straps 2 0.057 7800 7.6 @ 4000 lb

Lap belt 2-1/4 0.057 8880 7.8 @ 4000 lb

Lap belt

tiedown -

strap 1-3/4 0.057 6980 6.9 @ 3000 lb

*Based on 10-in. gage length.

150

S. . .. . .. _ _- . - . -,
I%II

Dynamic testing of polyester webbing has demonstrated the dy- . namic elongation to be approximately 60 to 75 percent of the

static elongation under the same load, as illustrated in Fig- ure 52 (References 63 and 83).

7.4.2 Width and Thickness Requirements

Selection of the optimum webbing width for a lap belt and

shoulder harness must be based on two conflicting requirements: * - (1) maximum width for lowest pressure and (2) minimum width

for maximum comfort and minimum hardware weight. The widths

* specified in Table 6 are believed to be a good compromise between these conflir[ing requirements. All webbing used for

restraint harnebs,. r st be thick enough to ensure that the

webbing does not'.A.. or crease to form a "rope" or present a

thin sharp edge under high loading that will cause damage to

soft tissue. Such damage is more likely to occur in the neck

region during a lateral loading or, in the pelvic region during

a forward loading. Although requirements based on early investigations using nylon webbing specified a minimum thickness of

0.090 in., it has since been determined that state-of-the-art

webbing materials must be thinner in order to achieve the desired low elongation. No significant problem of injuries

caused by the thin webbing has been observed with this lowelongation webbing which has seen extensive automotive use.

Therefore, based on currently available materials, a minimum

thickness of 0.055 in. is considered acceptable.

7.4.3 Webbing Attachment Methods

7.4,3.1 Stitched Joints: The strength and reliability of

stitched seams must be ensured by using the best known cord

sizes and stitch patterns for a specified webbing type. The

stitch patterns and cord sizes used in existing high-strength

military restraint webbings appear to provide good performance. The basic stitch pattern used in these harnesses is a "W-W"

configuration for single-lapped joints. Research by the U. S.

Naval Aerospace Recovery Facility (NARF) at El Centro, California, has reaffirmed the adequacy of basic "W-W" stitch patterns; however, the research also revealed that a larger--size

cord with fewer stitches per inch gave superior performance to

the No. 4 MIL-T-7807 cord currently being used. On the basis

of this research, the 50-lb strength No. 6 ccrd at 4-1/2 to 5

83. Kourouklis, G., Glancy, J. L., and Desjardins, S. P., THE

DESIGN, DEVELOPMENT, AND TESTING OF AN AIRCRAFT RESTRAINT

SYSTEM FOR ARMY AIRCRAFT, Dynamic Science, Division of Ultrasystems, Inc.; USAAMRDL Technical Report 72-26, Eustis

Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, June 1972,

AD 746631. 151
8000

7250

700 -

000Failure

000/ 6100 .0 Dynamic,'

5000 /

4000 -

S/ /I_,

100

0 5 10 15

Elongation, 04percent

Figure 52. Load elongation characteristics for

MIL-W-25361 (Type II) polyester I

webbing for static and rapid loading

rates.

TABLE 6. MINIMUM WEBBING WIDTH REQUIREMENTS

Minimum width

Webbing identity Ain.)

Lap belt 2-1/4'

Shoulder strap 2.0

Tiedown strap 1.5

*A greater width (up to 4 in.) or pad is desirable in the center abdominal area.

152

- ..- .. . . . ., . "
stitches per inch is recommended, as illustrated in Figure 53,

for use on MIL-W-25361 webbings. Also, the heavier cord can

be expected to provide better resistance to sunlight degradation and abrasion. The use of the 50-lb cord and an 80-percent

efficiency results in a minimum strength of 160 lb/in. (4

stitches x 50 lb/stitch x 80 percent) for a single-lapped joint

or 320 lb/in, for a looped joint. Thus, the total stitch

length needed can be determined by the total required load.

4-1/2 to 5 stitches per inch

MIL-T-7807B, No. 6 nylon cord

(Type I or 1I, Class I, 50 ib) Metal or equivalent seam strength hardware with smaller diameter cord for

wiThe new thinner webbing.

0.12 in. - . 0 o25-in.

spacing

S4-1. 5-in. [- -- 8One stitch

minimum minimum

Figure 53. Stitch pattern and cord size.

It has been shown recently that the heavier thread is not compatible with the new low-elongation polyester webbing (Reference 84). For these webbings, a smaller diameter cord offers

the advantages of reduced webbing fiber damage and the ability

to be used with automatic sewing machines and is therefore acceptable.

The strength of stitched joints can be expected to decrease

with age because of normal weather exposure and because of the

normal dust and grit collection between the webbing surfaces.

The grit and dust can gradually abrade the cords over a period

of time. The use of a 30-percent increase in the total stitch

length required is recommended to offset the normal aging

84. Farris, L., HIGH STRENGTH STITCHING FOR AIRCRAFT PERSONNEL

RESTRAINT SYSTEMS, Pacific Scientific Co.; Proceedings,

1978 SAFE Symposium, Survival and Flight Equipment Association, Canoga Park, California, October 1978.

153
| ~I I

DI

strength decrease as well as the possible abrasion strength j

decrease. Covering the stitched joints with cloth to provide

wear protection for the cords is also recommended.

An example of establishing the total seam length is given:

Assume: A single-lapped joint, 50-lb cord strength, with a

4000-lb joint load.

Then, the minimum stitch strength is

(50)(4)(0.80) * 160 lb/in.

and the minimum seam length is

4000 4000 - 25 in.

160

Therefore, the total seam length is

(25)(1.3) - 32.5 in. I

The total seam length is achieved through placing many short

lengths in a rather small area. Several patterns have been

developed and tested; however, the W-W as described below is

still preferred. The size of the overlapped and stitched area

should be minimized to reduce weight, reduce the stiffened section of the webbing, and provide more room between fittings for

adjustment.

Unpublished data from comparative tests of five stitch patterns performed by NARF indicated better performance of two

new stitch patterns over the basic W-W pattern. The data from

this research are reported here with permission of NARF. -

The five stitch patterns tested are shown in Figure 54. These

patterns were sewn in Types XIII and XXII of MIL-W-4088 nylon

webbing used for parachutes. Three samples of each stitch pattern were tested. Table 7 shows the results of the first test

series. Because of the low number of total stitches, the results were inconclusive, and a second test series was performed. Patterns 2 and 5 were eliminated from the second

series. Table 8 shows the results of the second test series.

It relates the performance of the two stitch patterns, I and

154
- --.---_<, 

" ~~~Pattern (j-

-- ,-i

• - -,'

Pattern -I

<.---------------- "-"-- ------ I

5 Stitches/inch

_ HI

S ...... - .- - " 

Pattern I -• - - - II

2 -- - - - - - - - - - II

7 Stitches/inch

Pattern --

3

5 Stitches/inch

I-- - - - --.----- ---

Pattern 5 S

Li

L-- T•-- -- -- -- --- ~.w- -

5 Stitches/inch--3

Pattern I . " --- I

I . .I -- -" I

5 Stitches/inch

Figure 54. Stitch patterns tested.

'4

S' 155

- ..... • . . ............--- . ,_ --. ."... , •
S0 0 0.1E

In Ln in Q

3,0 n V I vi LA m EN

&A00 0 P) f N t0 r- 0 0 -' %D 0 % D In 0 n In In In .

fn0 0 O t0o 0 14 o0 t- '0 CID ~0 0 In M 01 n n mIE'

4'l

In A .o N 0N in 0

go 0 In In 'A In 0D E N

vd ýq 0~r4 ; to

A In 0 LA OD 0 o vn In O 0 %0 Go 0

4 In _0 -0 4 4

m0 v OD 0%

.0 n 0 n 0 Ln 0 41 4~ ~~~~ 0 0 0 * 04 0- 

0 4 C4 0 4 E r.0 

0 -0 00 co v v %0 r-4 In OD 0 H 4 in IV In IV 9 4 EN m r

OO)

in0 ) In i

V qn 41 0 .NA4 N 0

M4 cc w

: *

.4 4 4

0 .0 o w4 A).go 0 w4

IA .14 -.4

156
"TABLE 8. BREAKING STRENGTH OF STITCH PATTERNS (TEST SERIES TWO)

Webbin and stitch

Sample (a) (a) (a) (a) (a) (a)

1 4400 4410 4540(b) 6340 6420 6215

Breaking 2 4710 4740 5080 6480 6490 6060

strength (ib) 3 4820 4360 4870 7200 6500 6070

Average breaking

strength (ABS) (lb) 4643 4503 4830 6673 6470 6115

ASS/ABS for

pattern 3 1.03 1.00 1.07 1.03 1.00 0.945

Approximate

total stitches 260 270 270 260 270 270

ABS/stitch (lb) 17.86 16.68 17.89 25.67 23.96 22.65

ABS/stitch/ABS

for pattern 3 1.07 1.00 1.07 1.07 1.00 0.945

(a) A designates MIL-W-4088 Type XIII nylon webbing.

B designates MIL-W-4088 Type XXII nylon webbing.

Numerials 1, 3, and 4 designate stitch patterns as shown in

Figure 54.

(b) Jaw separation 20-in. minimum. All other tests at 2-in.

minimum.

4, to the performance of pattern 3, the W-W pattern, for the

two different types of webbing. Stitch patterns 1 and 4 exhibited better strength properties than pattern 3 (W-W) when

Type XIII webbing was used. Pattern 4 did not perform as well

when Type XXII webbing was used, while pattern 1 again indicated better strength characteristics than did pattern 3.

The W-W stitch pattern as shown-Ln Figure 53 is still recommended until more conclusive information on these or other

stitch patterns becomes available.

157

- '
Ii

7.4.3.2 Webbing Wrap Radius: The wrap radius is the radius

of the fitting over which the webbing is wrapped at buckles,

anchorages, and adjusters, as illustrated in Figure 55. Detailed information on just how small this radius can be before

the strength of the webbing is affected is not availablel however, the 0.062-in. minimum radius shown is based upon the

geometry of existing high-strength restraint harnesses. This

radius should be carried around the ends of the slot as shown

in Figure 55 to preclude edge cutting of webbing if the webbing should be loaded against the slot end.

A0

webbing

minimum

typical.

Wrap radius (0.062-in. minimum)

should be carried completely

around ends of slot as shown by

the heavy line above

DETAIL A

Figure 55. Wrap radius for webbing joints.

7.4.3.3 Hardware-to-Webbing Folds: A possible method of reducing fitting width at anchorage, buckle, or adjuster fittings is to fold the webbing as shown in Figure 56. This reduces

the weight and size of attachment fittings; however, it can

also cause premature webbing failure because of the compressive force applied by the top layer of webbing to the lower against

the fitting slot edge. If this technique is to be used, tests

to demonstrate integrity are recommended. Also, for configurations that require two load paths, such as lap belts, where an

adjuster cannot hold the required 4000-lb load, the webbing is

looped through a full-width slot which halves the load in each

strap. An adjuster is then included in one strap. Adjustment

requires that the webbing be freely drawn through the fitting,

a requirement that folded webbing cannot meet.

7.4.3.4 Surface Roughness of Fittings: A surface roughness

of no more than RMS-32 is recommended to prevent fraying of

the webbing due to frequency of movement over the metal.

158

'I

|H fi"iii '' ii' '°i i' l ii' i' ~i i'~i'F':• '- j .... _a,,-, ' -. . .
F

0

-El

Figure 56. Webbing fold at metal hardware attachment.

7.4.4 Energy-Absorbing Webbing

Energy-absorbing restraint system webbing has been considered

for limiting loads on the occupant. The potential advantages

of energy-absorbing webbing are: (1) reduction of maximum load

exerted by the webbing on the occupant and (2) reduction of the

amount of elastic energy stored in the webbing. Webbings of

this type have been developed and are described briefly here

for information purposes. They are not recommended for use in

seating systems for the reasons presented below.

The principle of energy absorption for the first webbing material depends on a core wrap of fiberglass that breaks at a design loadl then, the outer cover of nylon wrap takes over the

loading, gripping the fiberglass until it breaks again. The

construction of the webbing varies, depending on the type of

force-versus-percent-of-elongation curve desired. For this

webbing, the general shape for the force-versus-elongation

curve includes a linear elastic region followed by a region of

constant force.

The construction of the second type of e iergy-absorbing webbing differs greatly from the first. It is made of polyester,

159

- Ii- _______________________________________________ ~ - - - --. --- C

r . . .. ..... _ - j - . . . .
• I I I I I I I I I I I |

and the energy absorption is produced by the filaments themselves. The polyester filaments are heat shrunk from their I

original sizes; and they do not return to the shrunk dimensions after the load application. This has the effect of

plastic deformation; and this property provides the energyabsorption capability of the material. The general shape for

the force-versus-elongation curve for this webbing is a constant rate in pounds per inch which makes inefficient use of

stroke distance.

A third type of energy-absorbing webbing material has been

evaluated for parachute applications at the U. S. Naval Aerospace Recovery Facility. The material is made by stitching

together two pieces of webbing. The two pieces of webbing

separate (peel) at a constant load by breaking the stitches

holding them together. The constant breaking force can be

varied by increasing or decreasing the number of stitches. Because of its construction, the material does not appear to be

suitable for use in aircrew restraint systems.

Because of other considerations, including primarily the increased potential for secondary impacts of occupants, energyabsorbing webbing is not recommended for use in seating systems.

The limited room available in aircraft requires that the strike

envelope be minimized. Therefore, the use of the lowest elongation available is specified.

7.5 RESTRAINT SYSTEM HARDWARE

7.5.1 General

The restraint system configured for use in a particular location in an aircraft will include various hardware selected on

the basis of a trade-off among such factors as crashworthiness,

weight, and cost. An aircrew system meeting the requirements

of MIL-S-58095(AV) that has been developed is illustrated in

Figure 57. The system shown in Figure 58, which is defined by 1

a draft military specification (Reference 85), offers improved

protection but is heavier and more expensive. For example, it

includes two inertia reels for the reflected shouluer strap

system, which reduces both lateral and forward motion. Its

use may be warranted where space is a problem and strike envelopes need to be minimized. Also, this system's use of lap belt

retractors rather than adjusters provides greater convenience

in ingress, greater comfort by eliminating the adjuster, and

greater crash safety by eliminating slack (preload held on the

lap belt by torsional spring in retractor). The weight of the

85. Proposed Draft Military Specification,

RESTRAINT SYSTEM, AIRCREW, September 1974.

160

I

S • ... ... w. _- .
Inertia reel

Inertia reel strap

'-l- -. 75-in. 'ide

webbing

2.00-in. wide webbing

Shoulder k I,

harness Adj uster

2.00-in. wide

webbing !'• • •...Lower shoulder

Lap belt-/ N • • Lap belt•

2.25-in. wi e-- Ii,:, aho

webbing Fixed fitting on

tiedown strap

• Tiedown strap

'•-1.75-in, wide webbing

Tiedown anchor

Figure 57. Aircrew restraint system.

system shown in Figure 57 is 5.50 lb and that of the system in

Figure 58, 8.50 1b, with the difference due mostly to the additional inertia reel and the two lap belt retractors of the

latter system.

The various hardware components involved in a state-of-the-art

restraint system are described below. information on production items is included where available.

161

-I - --- - --
1.75-in, wide Inertia reel webbing

Reflected

10, .,Astrap

Ref lected strap

Roller fitting

'A djuster

shoulder 5 2.00-in, wide

strapwebbing

!,I Plug-in fitting

Rtary buckle

Rtator .4vi Fixed fitting

wide

Tiedown strap webbing

1.75-in. wide webbing

~4-Tiedown anchor

Figure 58. Reflected shoulder strap restraint

system. (From Reference 85)

162
7.5.2 Buckles and Emergency Release

The buckle is of the single-point release type for all systems and provides positive release of all strap fittings (with the

exception of the one to which it is permanently attached).

These capabilities should help prevent entrapment of a wounded

occupant.

7.5.2.1 Aircrew Restraint Buckle: To facilitate egress in emergencies, a rotary-release buckle provides the advantage of

operation by a force applied in many directions. In one existing buckle, each fitting can be inserted and locked separately.

When the release handle is rotated, springs move the fittings

far enough so that none will reengage when the handle is released. This is an example of a desirable feature that will

protect against a potential hazard created by a fitting relocking upon release of the handle. For example, if one lap belt

fitting relocks, it could partially restrain the occupant as he

attempts an emergency egress.

7.5.2.2 Troop/Passenger Restraint nuckle: The restraint systems recommend d for troop seat insta-llations, as shown in Figure 45, include a single-point, lift-lever release buckle that

is permanently attached to one of the lap belt straps. The

lift-lever release mechanism replaces the rotary release here

for the convenience of troops or passengers who, because of

infrequent system use, might find it easier to use in emergencies since it resembles automotive hardware (References 63, 78,

and 86). The design of such a buckle is described in Reference 78.

7.5.3 Adjustment Hardware

The lap belt-length adjusters are located either at the center

of the lap belt near the attachment-release buckle or at the

side of the hips of the occupant below the iliac crests of the

pelvis. Shoulder strap adjusters are located as low on the

chest area as possible to avoid a concentrated pressure over

the collarbones of the seat occupant. It is possible for the

seat occupant to make strap adjustments easily -with either

hand. A downward pull on-the free end of the shoulder harness

straps tightens the shoulder harness, and an upward pull on the

* free end of the lap belt straps tightens the lap belt.

86. Reilly, M. J., ENERGY ATTENUATING TROOP SEAT DEVELOPMENT,

The Boeing Vertol Company; Report NADC-AC-7105 with Addendum NADC-73121-40, U. S. Naval Air Development Center,

Aerospace Crew Equipment Department, Warminster, Pennsylvania, May 1971.

163
7.5.4 Inertia Reels, Control, and Installation

7.5.4.1 Inertia Reels and Controls: Inertia reels currently

installed on the crewseats of U. S. Army aircraft are designed in accordance with the requirements in Reference 87. The design requirements specified in MIL-R-8236 are compatible with the

other restraint harness requirements listed in this chapter,

and it is recommended that the use of this specification be

continued.

Some discussion of inertia reel function is included here to

explain the two basic types of reels listed in MIL-R-8236. The

basic function of the inertia reel is to give the crewmember

full freedom of movement during normal operating conditions

while automatically locking the shoulder harness during an

abrupt deceleration.

The freedom of movemei'... is obtained by spring-loading the reel

cable or webbing to which the shoulder straps are attached.

.This allows the shoulder harnes to be extended without apparent restraint of the shoulders (only 6 lb at maximum exten- sion). The reel will be constantly taking up any slack.

There are two basic types of MIL-R-8236 reels. The impact - sensitive type requires a 2- to 3-G impact on the rtia reel

housing itself to lock automatically. Normal flight loads,

including severe turbulence, will not activate this reel.

The rate-of-extension type reel, although mechanically different, serves the same purpose. Its automatic operation depends

on the rate at which the inertia reel strap is reeled off,

which makes it a function of the rate of upper torso displacement away from the seat back, regardless of direction. The automatic operation of this reel can be checked at any time by a

jerk on the shoulder straps. The shoulder harness, after being1

locked automatically, reels up the slack in the strap every time the occupant bounces back toward the seat back. Eventually, the- occupant will find himself fftmly locked against

the seat back.

A few words about the manual control lever ,ill clear possible

misconceptions about its use. Both types ot reels have identical control levers, usually mounted under the seat pan, on the seat side, or at some other convenient location. The lever I

has two positions: manual and automatic. The manual position

87. Military Specification, MIL-R-8236, REEL, SHOULDER HARNESS, INERTIA LOCK, Department of Defense, Washington,

D. C.

164

S. ..... .. .... ..4

.4
permits the pilot to lock the reel if rough conditions are anticipated, or at any other time warranted. Normally, the contLol lever should be in the a'itomatic position to allow the

wearer to lean forward easily and reach all controls without

first having to release the control lever. MIL-R-8236 requires

that both reel types lock automatically before the occupant

travels more than 0.5 in. during an emergency deceleration.

In &ddition to the MIL-R-8236-type reel, which has the function

of preventing further strap extension, there are power-haulback

reels, which rapidly retract slack to apply a tensile load to

the belt, Generally, these systems, some of which use a basic

MIL-R-8236 inertia reel, are powered by a gas generator and

must be manually actuated prior to impact. Automatic actuation

by an acceleration sensor is not recommended because human tolerance considerations limit the haul-back velocity. By the

time the crash could be sensed, there would not be time to complete the haulback within tolerable accelerative limits.

7.5.4.2 Inertia Reel Installation in Rotary- and Fixed-Wing

Aircraft: Accident statistics indicate that rotary-wing aircraft frequently impact on their sides, or impact vertically with little longitudinal de- -ration. Therefore, it is concluded that all rotary-wir . VTOL aircraft should incorporate the rate-of-extensior reel, because a unidirectional

(-G ) acceleration (needec -o actuate the impact type reel)

might not :e present in all rotary-wing or VTOL aircraft accidents.

On the other hand, the study of about 92 fixed-wing aircraft

accidents, described in Volume II, revealed that only one accident occurred in which no longitudinal (-G ) acceleration was

present. On this basis, it is concluded tiiat a unidirectional-

(impact) type reel would be adequate for fixed-wing aircraft.

However, it is recommended that the rate-of-extension type reel

be used on all aircraft types to assure locking regardless of

load direction.

The inertia reel may be anchored to the seat back structure or

to the basic aircraft structure. The shoulder straps must be

maintained at the correct angle with respect to the wearer's

shoulder at all times as described in Section 7.3.4. If an

anchorage to basic structure is used, consideration must be

given to the possible seat bucket motion so that the shoulder

strap angle or length does not change by a significant amount

during energy-absorbing stroke. The reel should be mounted

and the webbing routed so that the webbing does not bear on

the reel housing. Excessive webbing loading of the housing

can produce housing and/or webbing failure as the housing is

not designed as a contact surface for loaded webbing.

t 165
8. SEAT STRENGTH AND DEFORMATION REQUIREMENTS

8.1 INTRODUCTION

Previous sections of this volume have presented background in- - formation to aid in understanding the problems involved in designing crashworthy seats and restraint systems. This chapter presents specific design and test requirements for seat systems

and litter systems. Occupant sizes and weights to be used in

the design are defined, as are the required static design strength-deformation relationships. Static tests to demonstrate the adequacy of the system in all loading directions

are presented. Finally, dynamic test requirements, to demonstrate that the seat systems, restraint systems, and litter

systems will provide the degree of protection desired, ar.; also

defined. Successful completion of all static tests and dynamic

tests are required to demonstrate acceptability of a design.

In this chapter, the direction of applied loads are referred

to in terms of forward or aftward, lateral or vertical, and

upward or downward. These terms, together with aircraft and

occupant axes, are defined in Chapter 2, Definitions, and refer

to seat loading in directions consistent with the aircraft

coordinate system. Thus, a forward load on a forward-facing

seat is in the positive x direction with respect to both the

seat and the aircraft. If the seat is a side-facing seat, the

forward load would be applied to the seat in the plus-or-minus

y direction, depending on whether the seat faces right or left

respectively in the aircraft. For an aft-facing seat, the forward load would be applied in the negative (-x) direction (towards the back of the seat).

8.2 RECOMMENDED OCCUPANT WEIGHTS FOR SEAT DESIGN

It is recommended that the upper and lower limits of occupant weights to be considered in seat design be based on the 95th

and 5th percentiles. Ideally, seat stroke limits should be

sized for the 95th-percentile occupant while the occupant acceleration limits should be determined for the 5th percentile. -

If this were done, the resistive forces would be tolerable

while the stroke lengths would also be adequate for all occupants in the design range (5th through 95th percentile). However, in most situations sufficient stroke distance will not

be made available in the aircraft to permit using the ideal approach; therefore, compromises will have to be made. Specific criteria for these cases are presented in this chapter and, for

the present, should be viewed as minimum state-of-the-art design goals.

166

.- I - -. ---- . . ~ . .-- --- --- - --- ~ - - -. . .- - -i
4

8.2.1 Crewseats

The design weight should be based on the typical weight of the

seated occupant, not the extremes. Although the weight of a

95th-percentile, combat-equipped aviator can be as high as

250 lb, it is believed that a majority of the flight hours logged in Army aircraft over the past 20 years have been noncombat

hours. Consequently, it is probable that crewmembers will be

lightly equipped. The restrictions placed on crewseats, including stroke length, control access, and seat armor, limit

the flexibility of design options. If the crewseats were to

be designed to protect occupants over the full range of weights

(144 to 250 lb), a weight-sensitive energy-absorbing system would be mandatory. Thus, the typical aviator weight recommended for crewseat design should not include combat gear.

Based on data of Reference 88, typical aviator weights are presented in Table 9.

TABLE 9. TYPICAL AVIATOR WEIGHTS

95th- 50th- 5thpercentile percentile pe,:zntile

weight weight weight

Item (lb) (lb)

Aviator

(Reference 88) 211.7 170.5 133.4

Clothing 3.1 3.1 3.1

Helmet 3.4 3.4 3.4

Boots 4.1 4.1 4.1

Total

weight 222.3 181.1 144.0

Vertical -

-effective

weight 175.2 142.3 112.6

88. Churchill, E., et al., AZ1

THROPOMETRY OF U. S. ARMY AVIATORS 1970, Anth'ropology Research Project; USANL Technical

Report 72-52-CE, U. S. Army Natick Laboratories, Natick,

Massachusetts, December 1971, AD 743528.

167

(it _ _
I

Variable-load energy-absorbing systems are highly desirable to

maximize efficiency and thus, protection in limited space.

Therefore, they should be incorporated in seat designs whenever possible.

8.2.2 Troop and Gunner Seats

The same percentile range of occupant sizes should be considered for troop and gunner seat designs. A greater variation

of clothing and equipment is used by troops than by aviators;

troop seats should be designed to accommodate them. The 95thpercentile occupant should be considered heavily clothed and

equipped, while the 5th-percentile occupant should be considered lightly clothed and equipped. Based on data contained in

References 31, 32, 63, 79, and 89, the typical weights of

seated troops in aircraft are shown in Table 10.

8.3 STRENGTH AND DEFORMATION

8.3.1 Forward Loads

In Section 4.7, it was shown that for a load-limited system

there is a minimum displacement that must be achieved if the

system is to remain in place during a given deceleration pulse.

Actually, all systems are load limited, although not necessarily through original intent. The inherent load-deflection

curve for any system imposes a definite limit on the system's

ability to resist impulsive loading. The objective of intentionally load-limited seat systems is to make the best use of

the space available for relative displacement of the seat and

occupant with respect to the airframe, while maintaining loads

on the occupant consistent with the type of restraint system

used and the occupant's capacity to survive the loads imposed.

The basic data used in developing the seat design curves presented in Figure 59 were obtained through a computer simula- tion of the seat/occupant system (see Reference 25) and from

the results of static and dynamic seat tests (References 31,

32, 63, and 79) using body blocks and eanthropomorphic dummies,

respectively. The computer simulation allowed the calculation

of the seat displacement for given load-limiting values. The

simulator included a realistic kinematic behavior of the occupant and the nonlinear effects of the restraint system. It is

estimated that the requirements given in Figure 59 are not conservative for the input pulses selected for design purposes.

89. U. S. Army, THE BODY SIZE OF SOLDIERS - U. S. ARMY ANTHROPOMETRY - 1966, USANL Te-hnical Report 72-51-CE, U. S.

Army Natick Laboratories, Natick, Massachusetts, December

1971, AD 743465.

168
TABLE 10. TROOP AND GUNNER WEIGHTS

95th- 50th- 5thpercentile percentile percentile

weight weight weight

Item (lb) (ib) (Ib)

Troop/Gunner

weight

(Reference 89) 201.9 156.3 126.3

Clothing

(less boots) 3.0 3.0 3.0

Boots 4.0 4.0 4.0

Equipment 33.3 33.3 33.3

Total

weight 242.2 196.6 166.6

Vertical

effective

weight

clothed 163.9 127.4 103.4

Vertical

effective

weight

equipped 197.2 160.7 136.7

These are a 30-G peak triangular pulse of 50-ft/sec velocity

change in the cockpit and a 24-G peak with 50-ft/sec velocity

change in the cabin area.

The static loads that the seat must withstand are obtained by

multiplying the load factors (G) shown in Figure 59 by the sum

of the total weight of the 95thzpercentile crewmember or passenger plus the weight of the seat and any armor or equipment

attached to or carried in the seat. For crewseats, the weight

of combat gear is not included (see Section 8.2.1).

Longitudinal displacement of approximately 6 in. for cockpit

seats and 12 in. for cabin seats measured at the seat reference point (the seat reference point may be projected to the

outside of the seat pan for measurement convenience) is the

practical limit for seats in existing Army aircraft. Since

169

6L1

* __ .~ . .- 
4 45 FW I

40 -Acceptable failure FD;

A area

~ "i AceptabExaplerformac

ý Ccki set fowad-acn set

30

a /SamlAcceptabes faiur areace,

W 25

~o 20 Base c___

S10

,:4~ ~Cockit seats

0 ..... A ~ A / ~ Ž

mcepasured fatluseatreaeec on

:1.7

25- . - - _ __ _ _ _ _ _ __ _ _- -.-
there is typically more room available in cabins than in cockpits, the advantages o! longer energy-absorbing strokes can

usually be achieved. Longer strokes permit the absorption of

equivalent energy a& lower loads and thus can serve to z Auce seat weight and increase the level of protection offered over

a wider occupant weight range.

In viewing Figure 59, it can be seen that for cabin seats 12 in.

of stroke enables the minimum limit load to be reduced to 15 G;

whereas, for cockpit seats a 20-G minimum limit load is required with only 6 in. of stroke. The 15-G and 20-G minimum

limit loads fix the G levels of the base curves for the cabin

and cockpit seat respectively. The available stroke will be

unique for each specific aircraft, and the energy-absorbing

mechanisms in the seats should be compatible with the available

stroke distances. If forward or sideward motion threatens to

limit the effectiveness of the vertical energy attenuating system or increase the possibility of severe injury caused by secondary impact of the occupant with items in the aircraft, then

energy-absorbing stroke in directions other than vertical

should not be used. The 6 in. and 12 in. allowed by the curves

of Figure 59 should be viewed as maximum distances which are subject to limitations of available space in each specific aircraft and location in the aircraft.

The initial slope of the cockpit seat base curve to 0.75 in. of deflection allows for elastic deformation consistent with a

relatively rigid crewseat while the lighter weight and more 4

flexible troop/gunner seat requires a lesser slope. The 30-G

and 35-G upper cutoffs reflect consideration of human tolerance 

tions, 

limits, 

and practical 

load variations 

limitations 

between 

of seat 

cockpit 

weight 

and 

and 

cabin 

excessive

loca- I

airframe loading.

8.3.2 Use of Design Curves

To be acceptable, a seat design must have a characteristic

load-deflection curve that rises to the left and above the

base curves of Figure 59 and extends into the region beyond

the upper curve. This discussion also appliesýto the lateral

strength and deformation requirements discussed in Section

8.3.6. In Figure 59, curves A, C, and E are acceptable curves,

but curve B is unacceptable because it does not reach the required ultimate strength. Curve D reveals inefficient use of

seat deflection by intruding into the base area. The seat is

deflecting at too low a load, thus absorbing less energy than

it could.

171

wi W ~
8.3.3 Aftward Loads

Large aftward loads seldom occur in fixed-wing aircraft accidents but may occur in rotary-wing accidents. A capability to

withstand 12 G is recommended for aftward loads for all seats.

This value will usually be automatically met by all seats meeting the forward load requirements. Occupant weight should be

the total weight of the 95th-percentile crewmember or trooper

as presented in Section 8.2.

8.3.4 Downward Loads

Human tolerance to vertical impact limits the acceptable forces

in the vertical direction for all aircraft seats. The maximum

allowable headward acceleration (parallel to the back tangent

line) for seated occupants is on the order of 23 G for durations up to approximately 0.006 sec. Since most back tangent

lines are oriented at a backward leaning angle of 13 degrees

from the vertical aircraft axis, tolerance to vertical impact

loads should be somewhat increased over the stated criteria.

In spite of this, however, the 48-G design pulse imposes the

requirement for energy absorption in the vertical direction by

some form of load limiting. The vertical dynamic response of

seat-occupant systems and, in particular, the effect of seat

behavior on the occupant deceleration excursions, has not been

sufficiently investigated to allow a full explanation of the

effects of this phenomenon. The factors affecting the response

of the seat and occupant and thus the final design of the loadlimiting system include:

"• Input pulse variables.

"* Orientation of the occupant and seat relative to the

resultant force vector.

"* Effective occupant weight.

"* Occupant spring rate and damping characteristics.

"* Weight of the movable part of the seat.

"* Spring rate and damping characteristics of the seat.

-e Spring rate and damping characteristics of the cushion.

* Available stroke distance.

172
* Force-deflection characteristic of the energyabsorption system.

* Any external influences such as those caused by loads

transmitted through dummy legs, or binding of the

seat mechanism.

Army-sponsored research in these areas is currently underway

and results will be incorporated into this document when available.

The effective weight in the vertical direction of a seated occupant is approximately 80 percent of the occupant's total

weight because the lower extremities are partially supported

by the floor. The effective occupant weight may be determined

by summing the following:

* Eighty percent of the occupant's body weight.

* Eighty percent of the weight of the occupant's clothing (less boots).

* One hundred percent of the weight of any equipment carried on the body above knee level. Combat gear

is not included in the effective weight of the pilot

or copilot (see Section 8.2.1).

The dynamic limit load for the load-limiting system should be

established by use of a load factor fGL) of 11.5. The dynamic

limit load is determined by multiplying the summation of the

effective weight of the seat occupant, and of the movable or

stroking portion of the seat, by 11.5. The resulting dynamic

limit load includes the total force resisting the vertical

movement of the seat in a crash; the dynamic limit load of the

energy-absorption system, simple friction, friction due to

binding, etc. This requirement is difficult to satisfy with

a sliding guidance system because the frictional load varies

with contact load which, in turn, varies with the impact load

vector direction. A relatively friction-free rolling mechanism

or collapsible structure is therefore recommended.

The 11.5-G design cftternon, taken from Reference 24 and modified to provide a tolerable deceleration of the 5th-percentile

occupant, considers the dynamic response of the seat and occupant. The factor of 11.5 was established to limit the decelerative loading on the seat/occupant system to less than 23 G

for durations in excess of 0.006 sec (the tolerable level for

humans as interpreted from the Eiband data) in crashes that do

not exhaust the stroke of the seat.

173

~i
Crewseats should be designed to stroke a minimum distance of

12 in. when the seat is in the lowest position of the adjustment range. This distance is needed to absorb the resi-dual

energy associated with the vertical design pulse. Further, the

load-limiting system should be designed to stroke through the

full distance available including the vertical adjustment distance. Since a vertical adjustment of ±2-1/2 in. from neutral

is typically required by crewseat specifications, proper design

can provide up to 17 in. of stroke, depending on seat adjustment position.

The minimum of 12 in. of stroke is recommended to provide the

minimum required level of protection. As illustrated later in

this section, even with 12 in. of stroke, heavier occupants in

more severe crashes will exhaust the available stroke distance

and bottom out. The following reasons point out the need for

obtaining the greatest possible energy-absorbing stroke from

the seat:

* It is easier to provide energy-absorbing stroke in

the seat than in the fuselage or landing gear. The

distance from the floor of the helicopter to the

ground is usually specified either directly or implicitly by overall dimensional requirements.

* The energy-absorption capacity of the seat is much

easier to demonstrate than that of the airframe, as

the energy-absorption capacity of the airframe is

difficult to predict and hardware is usually not

available for testing in the early design phases of

a new aircraft.

Energy absorption assigned to landing gear can be

lost depending on the type of terrain upon which

the aircraft crashes; i.e., soft versus hard, as in

marshes or water as opposed to a landing strip. Aircraft attitude at impact has a significant influence;

also a relatively high roll angle, for instance,

could render the landing gear energy-absorbing feature inoperative. *

Based on the two previous points, the seat is a lowrisk approach for providing energy-absorbing stroke.

* 12 in. of stroke has been shown to be practical in

existing aircraft.

If it is absolutely impossible to obtain a minimum of 12 in.

of stroke, a lesser amount is acceptable, but in no case should

it be less than 7 in. The reduced stroke is acceptable for a

174
retrofit application or for use in small aircraft in which it

is simply impossible to find the space for a 12-in. stroke.

In such cases a systems analysis is mandatory; the analysis

must show that occupant protection is equivalent to the system

in which the 12-in. stroke is available.

For retrofit applications, the maximum protection possible

should be obtained in any component being modified, i.e.,

* seats, gear, etc. Separate test criteria have been established

for seats not having the required 12 in. of stroke and are presented in Section 8.6.3.2 of this document.

Energy-absorbing systems should be designed for 11.5 plus 1 G

minus 0 G considering the effect of the dynamic loading rate.

To obtain the static test loads, dynamic limit loads should be

reduced by the amount due to rate sensitivity of the particular

device used. Further, in the design of the system the desired

total resistive load on the seat should be obtained by summing

the resistive load provided by the energy-absorbing system and

the resistive load resulting from friction and/or other mechanisms unique to the particular system. Thus, the resistive load

of the energy-absorbing subsystem must be reduced from the load

required to decelerate the seat by the amount of the other

stroke-resisting variables.

If the energy-absorbing system is to provide only one force

setting, the effective weight of the 50th-percentile occupant

from Tables 9 and 10 should be used for sizing it in order to

ensure a tolerable stroke for the majority of the occupants,

not exceeding the stroke limitations of the seat. These weights

are 142.3 and 160.7 lb for pilot/copilot and troop and gunner

seats, respectively.

The following is an example of the calculations made for a

seat designed to stroke under the decelerative load imposed by

a 50th-percentile crewmember. The average deceleration and

stroke of the 5th- and 95th-percentile seat occupants are ap- -

proximated. First, using weights from Table 9, the 50thpercentile effective weight is calculated according to

Wteff = 0.80 (Wt5 0 + Wtc) + Wth (35)

where Wteff x effective weight of 50th-percentile occupant, lb

Wt50 - nude weight of 5th-percentile occupant, lb

175

- 4-. ... I I I __I . . .I . .. . .I~ l f
Wt a weight of clothes, lb

Tc

Wth a weight of helmet, lb

Thus, Wteff - 0.80 (170.5 + 3.1) + 3.4

142.3 lb

which is shown in Table 9 as the effective weight of the 50thpercentile crewmember. The effective weights for the 95th- and

5th-percentile aviators are 175.2 and 112.6, respectively.

Assuming a 60-lb movable seat weight, the total weights that

the load-limiting system must be designed for are:

5th percentile: 172.6 lb

50th percentile: 202.3 lb

95th percentile: 235.2 lb

The 50th-percentile static limit load (LL) is calculated as

follows:

LL = GL Wteff = (11.5) (202.3) z 2326 lb

The static load factors for the 95th- and 5th-percentile aviators are then

G 2326

L95th 25.2 =9.9

2326 -

GL5th .6 13.5

Seat deceleration spikes approaching 23 G could be expected for

the seat occupied by a 5th-percentile aviator. Decelerations

of this magnitude would not be expected to cause severe injury.

Also, in more severe crashes, the stroke would exceed 12 in.

for a seat occupied by the heavier percentiles. This would

mean that the 95th-percentile vertical survivable crash could

not be fully protected against in the seat adjusted-down position. With the seat in the neutral or up position, however,

protection over the entire range might be provided (see Reference 24).

176
For comparison, the same type of calculations for a system

limit load sized for the 95th-percentile crewmember yields the

following:

LL = (11.5) (235.2) = 2705 lb

and

G 2705 GL50th = 270 13.4

2705 GL5th = = 15.7

The entire population in the 5th- to the 95th-percentile weight

range could be expected to receive deceleration spikes in excess of 23 G in seats in which the limit load was designed for

the 95th-percentile occupant. Also, the natural distribution

of occupant weights places the majority of aviator weights near

the 50th percentile. It is therefore expected that more overall protection can be provided by sizing limit loads for the

50th-percentile rather than for a heavier occupant.

In order to use the stroke distance available at maximum efficiency, regardless of occupant weight, a variable-force loadlimiting mechanism is desirable. With an infinitely variable

force system, the deceleration levels can be maintained within

acceptable limits (if the stroke is not exhausted) for the full

range of occupant weights for either crew or troop seats while

using equal stroke lengths for identical pulses. A compromise

is possible for a seat design that uses a load-limiting device

rather than collapsing structure. The device can be designed

to produce two or more limit loads that can be selected by the

seat occupant. An example of one such system is discussed in

Reference 24. The selection would be made on the basis of

aviator weight. For example, for a dual-limit-l-oad device,

"the lowest load device might be-established by using the weight

of a 5th-percentile occupant. The second force might be designed for the weight of a 50th-percentile aviator. In operation then, the aviator would be required to select a limit load

by movement of a lever or dial upon entering the seat.

It is recommended that at least a dual-level load limiter

(preferably three or more levels) be used to provide maximum

protection over the complete occupant weight range. As an

illustration, consider the limit-load factors calculated

on this page and on 181. With the limit load set for the 9 177
NI3

50th-percentile occupant weight, the calculated load factors I

were 9.9 for the 95th- and 13.5 for the 5th-percentile occupant

weights. This produces a negative variation of 1.6 G for the

heavy occupant and a positive variation :f 2 G for the lighter

occupant from the design factor of 11.5 G. If two load settings were possible, the variations could be halved producing

0.8-G and 1.0-G variations, respectively. An infinitely adjustable mechanism would reduce the variation to zero.

The interaction between the occupant and the movable seat

masses increases with seat mass. Therefore, the movable seat

mass should be minimized. For integrally armored seats, this i

mass is relatively high, varying between 60 and 120 lb. However, design in compliance with the criteria presented herein

should provide the desired degree of protection. "he criteria

were derived primarily from studies of an inter Iiy armored

seat prototype and projected to unarmored seats. The criteria

are expected to be somewhat conservative for seats of very

light (10 to 15 lb) movable sections.

Because troops do not have operational functions to perform and

troop seats are not armored, more flexibility exists in troop

seat design. Troop seats should be designed for the maximum

stroke feasible to maximize protection over the large weight

range represented by the fully equipped and lightly equipped

occupant. It is recommended that the full 17-in. seat pan

height normally considered desirable from the human engineering

standpoint be used for energy-absorbing stroke. It is further

recommended, as a minimum, that the limit load of the system be

sized using the 11.5-G load factor and the effective weight

of the 50th-percentile heavily equipped occupant (160.7 lb).

Variable-level load limiters sized as discussed previously are

also desirable for troop seats.

8.3.5 Upward Loads

A capabilitl to withstand a minimum upward load of 8 G is recommended for all aircraft seats. Occupant weight should be

that of the 95th-percentile crewmember or trooper as presented

in Section 8.2.

8.3.6 'ie.ral Loads and Deformation

The lateral load and deformation requirements for forward- and

aft-facing sents are presented in Figure 60. Two curves are

presented. One is for light fixed-wing aircraft and attack and

cargo helicopters, while the other is for other rotary-wing

aJ-craft. The deflections of the seat are to be measured by

recording the motions of the seat reference point. Occupant

weight should be as stated in Section 8.2 and should be that

of the 95th-percentile aircrew member or trooper.

178
25

Lateral

Acceptable failure area

M b•C (Example for forward-

"0 is Rejection .4 0 or aft-facing seat)

rea .Utility aed

area I j"e ,%C observat ion

SIhelicopters

13. T t e (y)ttack, ando

Trotal controlled deformation (y), in.

measured at seat reference point

Figure 60. Lateral seat load and deformation requirements

for all types of Army aircraft.

Lateral loading in the forward direction (aircraft reference

system) on side-facing seats should be the same as for forwardloading (Figure 59) except load limiting should be employed.

For crewseats, the lateral deflection should be minimized; however, it is doubtful if any great stiffness can be achieved in

lightweight hardware. As a matter of interest, many new armored buckets are made from Kevlar, a very tough and strong material in tension. Its resin-starved condition (required for

good ballistic protection properties) leaves it with a rather

low flexural modulus, particularly after the seat has had other

loa=s imposed. The material is also rather rate sensitive

tE 'f under high loading rates, soft under low rates). For

th. . eason, it is believed adequate, as a design goal, to attempt to limit the initial deflection to 1 in. with a 2-in. requirement. Because of the possible loading rate sensitivity

of the seat materials, it is considered acceptable to demonstrate compliance by analysis of test data. This analysis might include adjustment of the static test data by use of measured or known deflection and load data from dynamic tests.

179

-I
Further, in cases where wells are provided under the seats to

increase the available stroke distance, the deformation should

be elastic. This may allow the seat to realign itself with the

well prior to entry after the lateral and longitudinal loads

are relieved, as explained in Chapter 4.

8.4 OTHER SEATS

The requirements presented for crewseats and troop and gunner seats also apply to passenger seats and any other seat

installed in the aircraft for any purpose. Unique seats installed for special uses are not to be exempt.

8.5 PERSONNEL RESTRAINT HARNESS TESTING 1•

The restraint harnesses are to be statically and dynamically

tested along with the seat and/or structure to which they are

attached as noted in Chapter 7. However, the lap belt, shoulder straps, and tiedown straps, including all hardware in the

load path, should be statically tested separately to ensure

that all components possess adequate strength and to determine

elongation. The strength and elongation test requirements bf

restraint system subassemblies are specified in Table 4.

Specific component tests, including operational tests, are detailed in a draft military specification (Reference 85). However, all components and subassemblies should be statically load tested. Each subassembly should be tested to its full design load to demonstrate its adequacy. Elongation characteristics should be measured to document these data for comparison

with requirements and use in systems analyses.

8.6 STRUCTURAL TEST REQUIREMENTS

Both stati. and dynamic tests are recommended. Dynamic tests

of aircraft seats have shown that individual components capable

of maintaining the design loads often fail when tested in combination with other components. This could be a result of inaccurate analyses. However, it is recommended that all seat

and litter systems be tested as complete units. This is not

to imply that component tests are not useful. Component tests

can be extremely useful and should be used wherever possible

to verify required strengths. This practice is particularly

valid where finite-element analyses have boen used to accurately predict distribution of loads in redundant structures.

Upon acceptance of prototype systems tested under both static

and dynamic conditions, no further tests shotuld be required

except for quality assurance. Major structural design changes

in the basic seat system will require static retesting of the

180
new system to ensure that no loss in strength has been caused

. by the design changes. If the changes could affect the energyabsorbing, or stroking, performance of the seat, additional dynamic tests should also be conducted. Major structural design

changes are those changes involving principal load-carrying

members such as floor, bulkhead, or ceiling tiedown fittings,

structural links or assemblies, seat legs, or energy-absorbing

systems. Minor changes, such as in ancillary fittings, can be

accepted without a structural test. A significant weight increase, however, such as the addition of personnel or seat armor, would require additional testing. In summary, changes

that increase loading, decrease strength, produce significant

changes in load distribution, or affect the stroking mechanism

will require retesting.

All testing is to be conducted with the seat cushions in place

and, for seats with adjustments, the seats should be in the

full-up and full-aft positions unless another position is shown

to be more critical.

If desired, dynamic tests may be substituted for static tests;

however, loading in all principal directions are required. Alternate dynamic tests are presented in Section 8.6.1.9.

8.6.1 Static Test Requirements

8.6.1.1 General: The purpose of the static tests is to demonstrate that the seat has the strengths and other properties

required to provide the desired performance in all the principal loading directions. Static testing enables basic properties to be ascertained for known loads applied at a slow enough

rate so that seat response can be observed. Successful completion of the static tests does not guarantee passing the dynamic

tests, but it improves the chances. Weaknesses can be identified and corrected prior to conduct of the ultimate dynamic

tests. Also, due to the loading rate sensitivity of materials,

load distributions may be different in dynamic tests than they

are in static tests. Certain structures, statically soft, ma:

react as stiffer members under dynamic loading, and thus, pick

up more of the load than when the system was loaded statically.

Because of these reasons and because of dynamic avershoot, a

margin of safety has been added to the ultimate static load

factor on the design curves as compared to the peak accelerations of the dynamic design pulses. It is recommended that

this margin not be sacrificed for reduced weight.

Table 11 presents the static test requirements for complete seat units. The tests required include a series of unidirectional tests to determine basic seat strengths along the major

181

~...................................- .
TABLE 11. SEAT DESIGN AND STATIC TEST REQUIREMENTS

Percentile

Test Loading direction occupant Load/deformation

L ref. with respect to Load used in load dei

no. fuselage floor required determination reguirementsai

I Upward 8-G minimum 95 No requirement

n bd .+.0 G so See Section 2 Downwardb 11.5 -0 G 8.3.4

Aftward 12-G minimum 95 No requirement

4 Forward See Figure 59 95 See Figure 59

5 Combined

Forward e,f See Figure 59 95 See Figure 59

Downwardc 11.5 +2.0 G 50 Same as Test 2h

-1.0 G

f Lateral 9-G minimum 95 No requirements

6 Lateralg See Figure 60 95 See iigure 60

(a) ( ate aircraft floor or bulkhead should be deformed as detailed in

in Figures 61 and 62, simultaneously with, or prior to the conduct of all static tests and kept deformed throughout load application.

(b) If more than one load-limiter setting is provided, a represen- tative sample of settings spanning the range of loads should be

tested.

(c) If more than one load-limiter setting is provided, the highest

load should be used.

(d) Subsequent to the stroking of the vertical %,ergy-absorbing device, cockpit seats should carry a static load of 25 G, based on

the effective weight of the 95th-percentile clothed and equipped occupant per Section 8.2 plus seat without loss of attachment

to the basic structure except when the seat pan has stroked to

and is supported by the floor.

(e) In the event that no load-limiting device is used in the forward

direction, a 20-G load for cabin seats and a 25-G load for cockpit seats may be used for this combined loading.

(f) For seats employing vertical guides which could distort under

combined loading and cause binding, the maximum forward and lateral loads should be reached prior to initiation of stroking.

This sequence demonstrates whether the seat will stroke downward

after transverse loads-are applied.

(g) The lateral loads should be applied in the most critical direction. In the case of symmetrical seats, the loading direction

is optional.

(h) Failure to meet the ll.5-G +2.0/-I.0-G static vertical load

limit should not be cause for seat rejection if the seat vertical

energy-absorbing system meets dynamic load requirements. Wi) Plastic deformation is permissible: however, structural integrity must be maintained.

182
axes. A combined loading test is also required to evaluate

the seat performance under static conditions simulating the

most severe, unsymmetrical loading condition anticipated. All

static tests should be conducted under simultaneous conditions

of floor buckling and warping as illustrated in Figure 61 or

bulkhead warping as illustrated in Figure 62. The warping conditions must be introduced in the static test phase to evaluate

completely the performance of the seat under the most severe

requirements selected for design.

8.6.1.2 Unidirectional Tests: Where separate strength and

deformation requirements have been specified in Table 11 for

longitudinal, vertical, and lateral loading of seats, the loads

should be applied separately. Seats must demonstrate no loss

in structural integrity during these tests and should demonstrate acceptable energy-absorbing capacity.

8.6.1.3 Combined Loads: Seats must demonstrate no loss of

structural integrity under conditions of combined loading as

shown in Table 11 and should demonstrate ability to stroke in

the vertical direction with the transverse loads applied.

8.6.1.4 Load Application Method: The static test loads are to

be applied at the expected center-of-gravity location of the

occupant or occupants of each seat. The loads should be applied through a body block (see Section 8.6.1.5) restrained in

the seat with the restraint system. Figure 63 shows the location of the center of gravity that should be used as the initial static load application point for the seat occupant.

For the testing, the seat should be adjusted to its aftmost

and full-up position. The loads calculated by multiplying the

weight of the occupant and equipment plus the weight of the

seat by the required load factor should be applied continuously,

or in not more than 2-G increments while the load-deformation

performance of the seat is recorded. Maximum loads need not be

held for more than 1 sec. The maximum load reached, regardless

of duration, is to be used to assess compliance.

On integrally armored =rewseats, care should be taken to assure

that the loads are applied proportionally to the proper assembly or test item to simulate the loads that would typically be

carried by the restraint harness and the seat support structure.

"* In other words, the portion of the load that could be expected

to be restrained by the restraint harness should be applied to

the body block as described above. The portion of the load

representing inertial loading of the movable assembly should

be applied separately at the center of gravity of the appropriate substructure through another provision. For example, a

183

. . .. -_2. .. ... ... ... - . . .
F-,

Original 

*

position

VL20 position

0,

Figure 61. Suggested method of applying floor warping

for static testing of seats.

184

-- i
y :Hydraulic cylinder

~Test fixture structure

Test

structure

----De flected

' position

Seat .,inal

attachments ::•io

Bulkhead

Figure 62. Suggested method of applying bulkhead

warping for static testing of seats.

lever to proportion the load between the body block and movable

section of the seat, and a sling to apply the appropriate portion of the load to the bucket, can be used. For seats with

relatively heavy frames, the inertial load of the frame can be

applied separately at its appropriate center of gravity. This

technique, although adding complexity to the test setup, assures that all components in the seat and restraint system assembly have been tested to their approximate static design

loads and that, as far as a static test simulation can be extended, performance and structural adequacy have been demonstrated. For lightweight seats (less than approximately 45 lb

for total seat and restraint system), the total load can be imposed on the body block.

8.6.1.5 Static Load Body Block: The static test loads must

be applied through a body block contoured to approximate a

95th-percentile occupant seated in a normal flying attitude.

The body block must contain shoulders, neck, and upper legs,

and provide for passage of a belt tiedown strap between the

legs. The upper legs should be contoured to simulate the flattened and spread configuration of seated thighs and to allow

the proper location of the buckle. Critical pelvis dimensions

9 185

}J

..... - A *4,. .. ~.- * . .n

- -mm----_---_--._--
It I

1..i

15.0 in. Ur

Forward > Aftward 00 in.

5 vIe0.0 t

0.5 in.

3.0-in, radius

o m Seat reference ".r 3: point

ajo

Figure 63. Static load application point and critical

body block pelvis geometry.

are shown in Figure 63. Buttock contours must be provided

to permit proper fit in a contoured seat pan. The leg stubs

should be configured to permit proper seat pan loading as the

body block rotates forward under longitudinal loading; i.e.,

the leg stubs should be only long enough to provide a surface

to react the lap belt load. The side view of the buttocks

should include an up-curved surface forward of the ischial tuberosities to allow the forward rotation of the body block

while maintaining the primary contact between the ischial tuberosities and the seat pan through the cushions.

8.6.1.6 Deflection Measurements: Deflection should be measured as lose to the seat reference point as possible to eliminate seat structure rotational deformation from influencing

the test results. To simplify thesc measurements, the seat

reference point can be projected to the outside of the seat pan

or bucket.

186
Normally the restraint system will be attached to the seat.

However, if a unique situation should develop in which the only option for increasing crashwQrthiness is to attach the system

(lap belts and shoulder harness) to the basic aircraft structure rather than to the seat, certain factors should be considered. First, the forward and lateral deflection requirements of Figures 59 and 60 need not be considered because the

restraint harness limits torso and seat deflection. Second,

the vertical deflection of the seat pan still must be considered since the downward movement of the seat pan could cause

excessive slack in the restraint harness, or the harness could

limit the stroke ox the seat, depending on where the restraint

system is anchored. Neither of these conditions is acceptable

in the design.

8.6.1.7 Load Determination: The total load required for all

test directions, except vertical downward, is determined by

multiplying the required load factor from Table 11 by the

weight of the 95th-percentile clothed and equipped occupant

from Table 9 or 10 (Section 8.2) plus the weight of each

seat. The effective weight of the 50th-percentile occupant

should be used to calculate vertical components of loading

(Test Nos. 2 and 5 of Table 11) as discussed in Section 8.3.4;

the effective weight of the 95th-percentile clothed and equipped occupant should be used for the bottomed test (Test No.

2(d) in Table 11). The weight of that portion of the seat that

strokes with the load-limited portion of the seat must be added

to occupant weight to determine the total required load in the

vertical direction.

8.6.1.8 Multiple Seats: Multiple-occupancy seats should be

fully occupied when tested. If it is determined that the most

adverse loading condition occurs in other than full-occupancy

situations, additional tests should be run for those conditions.

8.6.1.9 Substitution of Dynamic For Static Tests: It is recommended that static tests be conducted because of the advantages previously described. In summary, static tests are more

economical to run than dynamic tests; because of their slow

rate of load application, closer real-time observation of seat

response to the loading is possibleaiand, static testing provides structural response information which is more comparable

to the static analyses typically used in the development of

present seat designs. In the future, when dynamic analysis

becomes more reliable, this latter point will no longer apply.

A significant consideration in static-versus-dynamic testing

is the cost of the hardware. Static testing can be conducted

with a minimum number of seats because the condition of the

seat can be monitored and judgments made as to its acceptability for continued testing. If failures due to previous tests

occur, parts can be replaced and the test economically rerun.

187
If final acceptance decisions are based on dynamic tests alone,

a considerably more rigorous dynamic test matrix is required to

enable testing in all the principal loading directions. Dy- namic tests are usually more expensive than static tests and

the increased number of tests will also require additional

hardware. If new hardware is not used for each test, the results may be inconclusive. If the seat passes the test, the

results are acceptable; but if the seat fails the test, another

test must be made since it will not be apparent whether the failure was due to damage inflicted during a previous test or

due to a basic design or manufacturing flaw.

If for any reason, dynamic tests are substituted for the static

tests previously described, then loading in all principal directions must be conducted. The dynamic test requirements are

presented in Figure 64. These three tests must be conducted

in addition to the two presented in Section 8.6.2, and all five

must be passed. These tests are to be conducted in accordance

with the same ground rules as those presented in Section 8.6.2

and are subject to the same testing parameters and evaluation

procedures. Further, the static upload of 8 G and the static

aftward loading of 12 G must be imposed and satisfactorily passed.

8.6.2 Dynamic Test Requirements

8.6.2.1 Dynamic Test Requirements for Seats Having at Least

12-in, of Vertical Stroke: All U. S. Army prototype seats

should be dynamically tested to the two conditions specified

in Figure 65. These test conditions were determined from the

design velocity changes presented in Volume II. Test 1 is required to ensure that the vertical load-limiting provisions will perform satisfactorily under simultaneous forward and lateral loading conditions. Test 2 is required to ensure that

the seat can resist the loads produced by the design pulse when applied simultaneously in the forward and lateral directions. A 50th-percentile anthropomorphic dummy complying with

the Code of Federal Regulations, Title 49, Part 572 specifieation for dummies (Reference 90) should be used to simulate the

seat-system occupant for Test 1. A 95th-percenti]e anthropo- morphic dummy simulating as closely as possible the features of

the 50th-percentile dummy described above should be used to

simulate the seat-system occupant for Test 2. Total weight including instruments of these two test dummies should be:

90. U. S. Code of Federal Regulations, Title 49, Chapter 5,

Part 572: ANTHROPOMORPHIC TEST DUMMY, Government Printing

Office, Washington, D. C., (Rev.) 1978.

188
VAX.-

4*9 m-in

_____ ~'' - ~ ... to

to4

load I

t sa 2I IS .5 04 .7

L~~ I

10 -

to in ft/s t 0303

2b croad&tc ac~e 0ockpi o.013 C.057 .03t 3

Dummy

i ia nert G min 0.03 1.03 105 10 load

r. Rx 19 19 19 19

AV in, ft/sec 42 42 425 25

n buvai nylcotr t see 0.06Ž 0.036 .061 .036 j tn ia 2 sec 0.100 0.100 .104 .124

lod

Gflmax 21 33 21 27

AV min, ft/sec 30 s0 30 30

Ligurt 64fiqureens-fwinmitsti

cfrr susttu statd tests

heicptra~ 1 sc0.5, 0.33189 03
-I m 0| m I I I I 

0

aa

- IL2

I I;

tO t•l ti

Time, sec

Cockpit lest* Cabin seats

Test Confiquration Parameter QuaLification R&D Qualification R&D

Dummy inertial

load

t sac 0.043 0.024 0.059 0.034

t2 ooc 0.061 0.061 0.087 o.087

Av min, ft/sac SO s0 50 So

10'

2 t I sec 0.066 0.038 0.081 0.046

St2 sac 0.100 0.100 0.127 0.127

G min 28 28 22 22

i nertial G max 33 33 27 27

load

Lv min, ft/sec 50 s0 o 0 50

Figure 65. Dynamic test requirements for qualification

and for research/development testing.

91i)
50th percentile

Pilot/Copilot = 181.1 lb

Troop/Gunner = 196.6 lb

95th percentile

Pilot/Copilot = 222.3 lb

Troop/Gunner = 242.2 lb

Dynamic testing of multiple occupant seats should be performed

with the maximum number of occupants specified for the test

seat. Additional tests should be run if it is determined that

the most adverse loading condition occurs in other than fulloccupancy situations. For both tests of Figure 65, adjustable

seats should be adjusted to the full-aft and up position of

the adjustment range. Plastic deformation of the seat is per- missible; however, structural integrity must be maintained in

all tests. For Test 1, the seat should limit the acceleration

as measured in the pelvis of the dummy to values which ensure

that the 50th-percentile clothed seat-system occupant (see Section 8.2) will not experience vertical, +G , accelerations in

excess of human tolerance as defined in SJctions 4.3 and 4.8

of Volume II (see Figure 24 of this volume). The roll direction (10 degrees right or left) for Test I should be the more

critical loading for the specific seat design.

When determining compliance of the achieved test pulse with the

dynamic Xest requirements of Figure 65 (or Figure 64, as described fn Section 8.6.1.9):

1. Determine the maximum acceleration and construct the

onset slope for the test pulse by the method explained in Section 8.6.3.

2. Compare the achieved onset and peak acceleration of

the test pulse with those allowed and presented in

Figure 65. The achieved onset slope should lie between the minimum and maximum onset slopes using the

values of t1 and t listed in Figure 65 for the specific test conditions. The maximum acceleration

should also fall between the upper and lower limits

allowed.

3. Integrate the actual acceleration/time curve of

the test pulse and establish the achieved velocity change. The velocity charie achieved should be equal

to or greater than that tabulated for the specific

test conditions.

191

A-_ -f[&
8.6.2.2 Special Dynamic Test Requirements for Seats Having

Less Than 12 in. of Vertical Stroke: In the event that the

application of a systems approach permits the seat to have

less than 12-in. minimum vertical stroke, additional requirements are made of the dynamic testing. First, it would be desirable to perform a full-scale crash test with the test specimen, including all assemblies involved in the energy-absorbing

process. This would include a section of the fuselage, landing gear, and the seat or seats. This approach is totally acceptable for demonstrating the dynamic response and acceptability of the system.

Since cost associated with the type of system testing described

above is usually prohibitive, a different approach is acceptable. This approach includes dynamically testing the seat only,

as is done for systems with at least 12 in. of stroke, but modifying the input pulse to represent the energy-absorbing processes of the gear and fuselage. An example of such a modified

test pulse is presented in Figure 66. The initial plateau (t 1

to t') represents the acceleration-time history created by

stroJng of the landing gear. The sharp increase in acceleration at t' relates to fuselage impact, and the pulse beyond t'

representR the crushing of the stiffer fuselage section. The

velocity change under the pulse should be the same as identified for the particular crash force direction for other established tests (50 ft/sec for Test No. I or No. 2 of Figure 65).

G2

0

4.J

 .. Fuselaq

Landing gear stroking crushing

t t0 ti

Time, sec

Figure 66. Example of input pulse for seats

having less than 12 in. of stroke.

2.92

[

. ___ - _ . - - . . . .
It will be difficult to determine accurate dynamic crush characteristics of the various portions of the system to enable

establishment of a representative, and thus acceptable, test

pulse. The best analytical techniques, supported by test data,

should be used for determining the properties of the fuselage.

Since drop tests of landing gear are required, a much more accurate approach exists for obtaining the landing gear influence

on the pulse. Seat testing should await completion of landing

gear tests so that the results can be used to establish the

initial plateau (or other shape) between t1 and t6 of the input

pulse.

Typically the landing gear will stroke at loads below those

required to stroke the seat; therefore, much of the kinetic energy of the occupant and seat will be absorbed prior to fuselage impact. If the systems analysis is accurate, the energyabsorbing capacity of the seat will be sufficient to absorb

the residual energy at limit loads tolerable to the occupant.

Since each system may display different characteristics, it is

not appropriate to present in this document specific quantitative limits for use in evaluating the acceptability of the test

pulse. However, the same general approach and tolerances already presented for the standard pulse apply and should be

used. The technique described in Section 8.6.2.1 for establishing compliance with the required test pulse applies directly to the portion of the special test pulse following t'.

8.6.3 Data Acquisition and Reduction

Data acquisition and reduction should comply with the requirements of SAE J211 (Reference 91) for measurements of an anthropomorphic dummy, body accelerations, and structures.

Data should be presented in both analog and tabular form in

compliance with the sign convention shown in Figure 4. Impact

velocity should be determined and recorded for the test platform or vehicle. In the analysis of the data, velocity change

should be computed through either electronic means or graphically with a planimeter by integrating the area under the measured acceleration-time trace.

91. SAE Recommended Practice, SAE J211b, INSTRUMENTATION FOR

IMPACT TESTS, SAE Handbook 1979, Part 2, Society of Automotive Engineers, Inc., Warrendale, Pennsylvania, 1979,

pp. 34.117-34.118.

193
The method recommended for use in establishing the acceptabilparameters 

ity of the 

associated 

pulse (see 

with 

Section 8.6.2) and to determine other

sented the data is similar to that pre- in MIL-S-9479(USAF); 

as rise time, onset 

see Reference 92. Parameters such

may be slope, and acceleration plateau duration obtained using the following graphic approximation nique as shown tech- in Figure 67.

Gp 0. 9 Gp 3

Calibration bseline

j0.1 Gp

to tr

Figure 67. Graphic approximation example. (From Reference 92)

9 Locate the calibration baseline.

* Determine the maximum (G ) acceleration magnitude.

* Construct a reference line parallel to the calibra- tion baseline at a magnitude equal to 10 percent of the peak acceleration 

tersections (G ). The first and last in- of this line-with the acceeration- time plot defines points I and 2.

92. Military 

EJECTION, Specification, MIL-S-9479, SEAT SYSTEM, UPWARD AIRCRAFT, GENERAL SPECIFICATION FOR, Department of Defense, Washington, D. C.

194

p- ~ - - -
' 1

• Construct a second reference line parallel to the *

calibration baseline at a magnitude equal to 90 percent of the peak acceleration. The first and last

intersections of this line with the acceleration-time

plot define points 3 and 4.

* Some logic and practical judgment may be required for

selection of the first and last intersections depend- ing on the noise, structural or electronic, apparent

in the data. Significant tendencies are important,

not noise.

* Construct the onset line defined by a straight line

through points I and 3.

9 If desired, construct the offset line defined by a

straight line through points 2 and 4.

* If desired, construct a line parallel to the calibration baseline, through the peak acceleration. The

time interval defined by the intersections of this

line with the constructed onset and offset lines

L (points 5 and 6) is the plateau duration (Ot).

* Locate the intersection of the constructed onset line 1, with the calibration baseline (point 7). The time

interval defined by points 7 and 5 is the rise time

(tR). Referring to Figure 65, the rise time should

beRgreater than t but less than t when determining

compliance with d4namic test requA~ements. Point 7

is the initial time to in Figure 65.

8.6.4 Seat Component Attachment

Since components that break free during a crash can become lethal missiles, it is recommended that attachment strengths be

consistent with those specified for ancillary equipment (see

Section 6.6.5.9, Volume III). Therefore, static-attachment

strengths for components, e.g., armored panels, should be as

follows:

Downward: 50 G

Upward: 10 G

Forward: 35 G

Aftward: 15 G

Lateral: 25 G

These criteria may be somewhat conservative for load-limited

seats; however, load limiting is mandatory in the vertical direction only. In light of the potential hazard, the strength

requirements are considered justified.

195

fr~puwr"i
9. LITTER STRENGTH AND DEFORMATION REQUIREMENTS

9.1 INTRODUCTION

This chapter presents strength and deformation requirements

for litter systems. Aircraft systems are rather difficult to

design because of limitations including that of the strength of existing litters and width of utility aircraft as compared

to the length of standard litters. The ultimate vertical

strength of existing litters with a 200-lb occupant and a total

system weight of 250 lb (see Section 9.2) is about 13 G. Since

the desired decelerative loads to be imposed on these litters

exceed 13 G, special techniques must be used to limit the deflection and to support some of the occupant load. A new

litter should be developed having the required strength to support loads in excess of 13 G, preferably 17 G, as presented as

a minimum in this chapter.

The other problem is associated with the length of the litter.

The width of the new Army utility helicopter does not allow

litters to be placed in the preferred lateral direction. The

lateral orientation is preferred because of the characteristics

of existing restraint systems used on litters which provide

more support when loaded laterally than when loaded longitudinally. Since higher loads are more frequently seen in the

forward direction than in the lateral, it would be desirable

to orient the litters laterally in the aircraft. This is not

possible because the helicopter is not wide enough, so special

devices have been developed to permit loading the litters in a

lateral direction and then rotating the litters into a foreand-aft orientation inside the aircraft. Improved litter restraint systems are required to provide the desired support to

the supine occupant on litters oriented in the fore-and-aft direction in these aircraft.

This chapter presents the design strength-deformation relationships and testing requirements for aircraft litters and their

supports.

9.2 RECOMMENDED OCCUPANT WEIGHTS FOR LITTER DESIGN

The litter strength and deformation requirements defined below

are based on a 200-lb, 95th-percentile litter occupant with

20 lb of clothing and personal gear, a 10-lb splint or cast,

and 20 lb of litter and support bracket weight for a total

weight of 250 lb (the weight of a litter and patient as specified in MIL-A-8865 (ASG), Reference 93).

93. Military Specification, MIL-A-8865, AIRPLANE STRENGTH AND

RIGIDITY MISCELLANEOUS LOADS, Department of Defense, Washington, D. C.

196
9.3 VERTICAL LOADS

9.3.1 Downward Loads

In the case of litter systems, human to&erance is not the lim- iting case in the vertical direction. The loads would be applied in a transverse direction to the body of a litter occupant. However, design to the 45-G human tolerance level is

impractical due to the strength requirements for litters and

for the basic structure to support the litter systems.

Litters are either hung from the ceiling or supported at the

floor. In either case, the input deceleration pulses are the

same as for floor- or bulkhead-mounted seats (see Volume II).

The use of ceiling-supported litters is limited by the strength

of the overhead fuselage structure. The inefficiency of structural deformation of the ceiling of older aircraft requires

additional energy-absorbing stroke to provide the protection

desired. Litters should not be suspended from the overhead

structure unless it is capable of sustaining, with minimum

deformation, the downward loads from the tiers of litters.

Therefore, in the design co an efficient system, intentional

load limiting should be related to the floor pulse.

The vertical strength and deformation requirements for a litter

system are detailed in Figure 68. This curve is read in the

identical manner as the seat load-deflection curve shown in

Figure 59. The load factors in units of G are based on the

summation of the weights of the occupant plus clothing, personal gear, splint or cast, and the weight of the litter and

attachment brackets for a total of 250 lb as described in Section 9.2. The curve of Figure 68 is based on the assumption

that 3 or 4 in. of vertical deflection will occur at the midpoint of the litter. In the unlikely event that a rigid litter

is used, an additional 2 in. of deflection should be added to

the curve. The deflection curve is limited to 6 in., because

a larger deflection occurring on one corner of the litter due

to an asymmetric loading could cause ejection of the litter occupant. A larger- energy-absorbing stroke can bd used effectively if a mechanism-is included in the system to control the °

amount of tilt allowed. For example, a system mechanism could

be designed that forced all four corners of the litter to

stroke the same distance (within elastic limits) thus achieving

this goal.

The additional problem associated with inadequate litter

strength must be dealt with in the design of litter systems.

The curve of Figure 68 assumes a litter capaile of at least

17 G with a maximum of 25 G. If the existing litter is used,

then a pan, net, or other device should be included under the

197

S..~. . . .

-v'" ~ ,- -
Aircraft ceiling

Downward fl j

Aircraft floor

ACCEPTABLE

0

m f : ::!: BASE AREA ;"l:::xi

10

0 1 2 3 4 5 6

Total controlled

deformation (z), in.

*G value based on 250-lb per litter

position.

Figure 68. Litter downward load and

deflection requirements.

litter to catch and support the litter occupant if the litter

fails. Actually the device should limit the deflection to a

value less than required to fail the litter and should stroke

with the litter. If all of these provisions are included, i.e.,

a rigid new litter or old litter with supporting pan underneath, together with the tilt-limiting mechanism, then the

stroke can be extended to 12 in. at a 17-G limit-load factor.

The load-deformation curve of Figure 68 would be extended at

17 G to 12 in. of stroke.

198

iA
Further background information on analysis and testing of helicopter litter systems can be found in Reference 22.

9.3.2 Upward Loads

All litter systems should be capable of withstanding a minimum

upward load of 8 G.

9.4 LATERAL AND LONGITUDINAL LOADS

Litter systems for all aircraft should be designed to withstand

the load and deformation requirements indicated in Figure 69

in all radials of the lateral/longitudinal plane. The litter

lateral loads are made equal to the longitudinal loads because

the litters may be oriented in either direction depending upon

the aircraft.

I ~ ~30. ..

Acceptable failure area

A ceotable

25

M

-4 20 Unacceptable

15

M

0 1

0 1 2 3 4 5 6

Total controlled forward** deformation of litter bed, in.

*G value based on 250-lb per litter position.

**Forward is the direction towards the nose of the aircraft

regardless of litter orientation in the aircraft.

Figure 69. Litter forward or lateral load and deflection

requirements for all types of Army aircraft.

4 199
The 20-G acceptable load level indicated in Figure 69 is predicated on the tolerance to acceleration of an individual restrained by straps on existing "table top" litters. If litters

and allied restraint harnesses are designed for improved crashworthiness, the 20-G load should be increased to 25 G.

Acceptable or nonacceptable load-deformation characteristics

are read from Figure 69 in the identical manner as the readings

from Figures 59 and 60 for seats. The deformation is measured

with respect to the aircraft floor along the longitudinal axis

toward the nose of the aircraft, regardless of litter orientation.

9.5 LITTER RESTRAINT HARNESS TESTING

The restraint used in existing military litters consists of two

straps wrapped around the litter. These straps should withstand a straight tensile minimum load of 2000 lb (4000-lb loop

strength). The maximum elongation should not be more than

3.0 in. under the straight pull (end-to-end) test on a minimum

strap length of 48 in. Elongation is restricted for litter

belts in order to minimize dynamic overshoot.

9.6 LITTER SYSTEM TEST REQUIREMENTS

9.6.1 Static Test Requirements

9.6.1.1 General: Table 12 presents the static test requirements for complete litter systems. Since previous studies have

shown that existing litters will not withstand the loads as

specified in this chapter, the assumption must be made that a

litter of sufficient strength will be developed prior to implementing these recommendations. The tests required include a

series of unidirectional tests to determine basic litter and

attachment strengths in the major axes. Also, a combined loading test is required to evaluate the litter system performance

under static conditions simulating a severe crash loading situation with loading components in multiple directions. Since

the litter orientation can bg either lateral or longitudinal,

a single requirement is made for-transverse loading in the horizontal plane (Test 5).

9.6.1.2 Unidirectional Tests: The test loads for forward, lateral, and downward loading of litter systems as presented in

Table 12 should be applied separately.

9.6.1.3 Combined Loads: Litter systems must demonstrate no

loss of system integrity under conditions of combined loads as

specified in Table 12.

200
igu

TABLE 12. LITTER SYSTEM STATIC TEST REQUIREMENTS

Test Loading direction

ref. with respect to Deformation

no. fuselage floor Load required requirements

1 Forward See Figure 69 See Figure 69

2 Lateral See Figure 69 See Figure 69

3 Downward See Figure 68 See Figure 68

4 Upward 8 G No requirement

5 Combined loading

Downward plus See Figure 68 See Figure 68

transverse load

along any radial

in the x, y plane

of the aircraft See Figure 69 See Figure 69

' 9.6.1.4 Point of Load Application: The loads should be ap- plied through a body block that simulates a supine occupant.

9.6.1.4.1 Forward (Longitudinal) - Lateral Tests: For systems

using the existing litter, a rigid simulated litter may be substituted for the actual litter. This will enable application

of equal loads at all attachment points between the litter and

the suspension system and allow testing of the suspension

system. The rigid litter substitution does not apply if the

litter has adequate strength to take the loads.

9.6.1.4.2 Downward and Upward Tests: Downward and upward

loads may be applied to each vertical suspension point separately. If the suspension system has the tilt-limiting -eatures, and the litter is adequate, then the load should be applied at the center of gravity of the body block.

9.6.1.5 Deflection Measurements: Downward, forward (longitudinal), and lateral deflections should be measured at the

bracket attaching the litter to the suspension system.

9.6.1.6 Load Determination: The test load should be determined by multiplying the required load factor (G) as specified

in Table 12 by 250 lb.

201

I

II
II

9.6.2 Litter System Dynamic Test Requirements

A single test to evaluate the vertical load-limiting system is

required. Litter systems with 95th-percentile anthropomorphic

dummies litter should and 30 be lb subjected (250-lb total) to a triangular of additional acceleration weight in pulse each

of 48-G peak and 0.054-sec duration (42-ft/sec velocity change).

The same test pulse tolerances, data, handling, and processing requirements as presented for the seats in Section 8.5 apply.

At least three accelerometers should be placed in the dummy;

one in the head, one in the chest, and one in the pelvic region. The instruments should be positioned to sense accelera- tions in the vertical directions (x axis of the supine occupant,

z direction relative to the aircraft). The input acceleration- I

time pulse also should be measured. It is advisable to use redundant accelerometers to sense the input pulse to assure acquisition of the needed impact environment data.

I!

, I

202t

kI

Li
IIIII II m

10. DELETHALIZATION OF COCKPIT AND CABIN INTERIORS

10.1 INTRODUCTION

The kinematics of body action associated with aircraft crash

impacts are quite violent, even in accidents of moderate severity. The flailing of body parts is much more pronounced when

the aircraft occupant is restrained in a seat with only a lap

bilt. However, even with a lap belt and a shoulder harness

that are drawn up tightly, multidirectional flailing of the

)-ad, arms, and legs, and to a lesser extent, the lateral displacement of the upper torso within its restraint harnessing,

is extensive. If it were possible to provide adequate space

within the occupant's immediate environment, this flailing action of a fully restrained occupant would not be a particular

problem. Since space for occupants is usually at a premium in

aircraft, especially in cockpit areas, it is not feasible to

remove structural parts of the aircraft sufficiently to keep

the occupant from striking them. The only alternative is to

design the occupant's immediate environment so that, when the

body parts do flail and contact rigid and semi-rigid structures, injury potential is minimized.

An occupant who is even momentarily debilitated by having his

head strike a sharp, unyielding structural object or by a leg

injury can easily be prevented from rapidly evacuating the aircraft and may not survive a postcrash fire or d water landing.

The importance of occupant environment designed for injury prevention, therefore, sbould be emphasized if optimum crash protection is to be ensured.

Several approaches are available to alleviate potential secondary impact problems. The most direct approach, which should

be taken if practical, is to relocate the hazardous structure

or object out of the occupant's reach. Such action is normally

subject to tradeoffs between safety and operational or human

engineering considerations. If relocation is not a viable alternative, the hazard might be reduced by mounting the offending structure on frangible or energy-absorbing supports and applying a padding material to distribute the contact force over

a larger area.

10.2 OCCUPANT STRIKE ENVELOPES

10.2.1 Full Restraint

Body extremity strike envelopes are presented in Figures 70

through 72 for a 95th-percentile Army aviator wearing a restraint system that meets the requirements of MIL-S-58095(AV)

(Reference 14). The restraint system consists of a lap belt,

203

t
10 i.n.

IHeel ret line for cockpit.

Air-c-ra t floor line for

troop compartment.

Figure 70. Full-restraint extremity strike

envelope - side view.

lap belt tiedown strap, and two shoulder straps. The forward

motion shown in Figures 70 and 71 was obtained from a test

utilizing a 95th-percentile anthropomorphic dummy subjected to

a spineward (-G ) acceleration of 30 G. The lateral mot-ion is-_

based on an exthapolation of datar7frorn the same 30-G test. In

positions where an occupant is expect-ed to wear a helmet, the

helmet dimensions must be added to the envelope of head motion.

204

7 . 4 . - _______________ - .-- --- ill" -.
enelp -- tovew

/ ___i__. /

Figure to 71.upr FUll-restraint- reqxtremity strike Ay i

10.2.2 Lap BeJ~t-Only Restraint

Although upper torso restraint is required in new Army air- 1•~

craft, strike envelopes for a 95th-percentile aviator wearing

lap belt-only restraint are presented in. Figures 73 through 75

for possible use. They are based on 4-G accelerations and

4 in. of torso movement away from the seat laterally and forward. In positions where an occupant is expected to wear a helmet, the helmet dimensions must be added to the envelope of head motion.

205

9••., - ..
0- -i-

/

I

t i in00

.. He.l rest line for cockpit.

Aircraft floor line for

troop compartment.

Figure 72. Full-restraint extremity strike

envelope - front view.

10.2.3 Seat Orientation

The strike envelopes of Figures 70 through 75 apply to all seat

orientations.

10.3 ENVIRONMENTAL HAZARDS

10.3.1 Primary Hazards

The primary environmental hazards are those rigid or semirigid

structural members within the extremity envelope of the head

and chest. It can be seen in Figures 70 through 75 that the

strike envelopes allow considerable upper torso movement for

various seating and restraint configurations. Since the upper

206
10 in.

* Heel rest line for cockpit.

oor line for

troop compartment.

Figure 73. Lap belt-only extremity strike

envelope - side view.

torso, and particularly the head, is the most vulnerable part

of the body, maximum protection must be provided within itsstrike envelope.-

10.3.2 Secondary-Hazards

Secondary environmental hazards are those that could result in

trapping or injuring the lower extremities to the extent that

one's ability to rapidly escape would be compromised. The

movement of unrestrained lower extremities in a crash impact is

not significantly influenced by method of body restraint. Consequently, even with an optimized body restraint system, those

areas within the lower extremity strike envelope must include

ample protective design.

207

strie enelop. I

10.32 Seonday Haard

Secodar haard envronenta arethoe tht culd esut i

trapingor he lwerextemites ijurng o te exenttha

ones biltyto apdl esap wul be- copoisd.h
-.. -. -% 

F

r'I 

-

- -

-9

K' -1 _ _ _ _ /

/ 

S

.140

4)S0S ______ ______ I. 00

4)

41S

4.)

'4S

p-I 0

41

'U

i-10

LI

.14

208
$4p

31c 0

.14 0

P-0 rq,

Lm0

I41

*14

209
10.3.3 Tertiary Hazards

Tertiary environmental hazards are those rigid and semirigid

structural members that could cause injury to flailing upper

limbs to an extent that could reduce an occupant's ability to

operate escape hatches or perform other essential tasks.

10.4 HEAD IMPACT HAZARDS

10.4.1 Geometry of Probable Head Impact Surfaces

Aircraft in the U. S. Army inventory in 1965 were examined to

determine the kinds of contact hazards most commonly found

(Reference 94). Typical hazards in the cockpit area included

window and door frames, consoles, control columns, seat backs,

electrical junction boxes, and instrument panels. Reference 94

presents further details of these impact hazards and a statistical analysis of head injuries in both civilian and military aircraft accidents. Contact hazards commonly found in aircraft

cabin areas include window and door frames, seats, and fuselage

structure. Use of suitable energy-absorbing padding materials,

frangible breakaway panels, smooth contoured surfaces, or ductile materials in the typical hazard areas mentioned will reduce the injury potential of occupied areas.

10.4.2 Tolerance to Head Impacts

Protection of the head in the form of protective helmets and

energy-absorbing structure and padding in the occupant's immediate environment is considered to be essential since, under

certain circumstances, even the forces incurred in minor crash

impacts could cause unacceptably high head impact velocities.

Tolerance levels for head impact are discussed in detail in

Volume II, and the reader should refer there for an understanding of the problem. However, for the case of forehead impact

on a flat surface, which is pertinent to the discussion of this

section, the most widely accepted collection of tolerance data

is represented in the tolerance curve of Figure 76. These data,

resulting from impact tests conducted on animals and human cadavers at Wayne State University, demonstrate the contribution

of both acceleration and pulse duration to the tolerance criterion.

94. Haley, J. L., Jr., et al., HELMET DESIGN CRITERIA FOR IMPROVED CRASH SURVIVAL, Aviation Safety Engineering and Research (AvSER), Division of Flight Safety Foundation,

Inc.i USAAVLABS Technical Report 65-44, U. S. Army Aviation Materiel Laboratories, Fort Eustis, Vircinia, January 1966, AD 628678.

210

i
600 5oo IT 7, 

Dengero s to life 300 _-__ 4--

'In

Time duration of effective

acceleratio~n, meec

brai dan in eheadags aintplane,

unyielding sufcs (Fo eeec 5

10..3 estProcedures

Thesimpitest procedure for evaluating the effectiveness

structure and padding in preventing serious head

Figur;e 400an SaeTlrac uv for the humann

injury makes use of an instrumented headform. The headform,

equipped with an accelerometer, can be propelled by a ram,

dropped, or swung on a pendulum to impact the surface to be

evaluated. The recommended procedure is described in SAE J921

(Reference 96). The measured acceleration pulse can be averaged for comparison with the Wayne State Tolerance Curve, or

integrated to compute a Severity Index, as discussed in Section 4.4.1 of Volume LI.

Figure 77 shows typical head velocities relative to the seat

as measured on anthropomorphic duammies, cadavers, and live

hurin subjects in dynamic seat tests. Various combinations

95. Patrick, L. M.1 Lissner, H. R., and Gurdjian, E. S., SURVIVAL BY DESIGN - HEAD PROTECTION, Proceedings, Seventh

Stapp Car Crash Conference, Society of Automotivengineers, Inc., New York, 1963.

96. SAE Recommended Practice, SAE J921b, MOTOR VEHICLE INSTRUMENT PANEL LABORATORY IMPACT TEST PROCEDURE - HEAD AREA,

SAE Handbook, 1979, Part 2, Society of Automotive Engineers, Inc., Warrendale, Pennsylvania, 1979, pp. 34.133-

34.134.

-L-1
- live volunteer

* " cadaver

S+- 95% dummy

lIbo - lap belt only

lbsh - lap belt and shoulder harness

60

4. 30 . .. .. ..

4 0

0 V__

o 20 40 60 80

i Seat velocity change in test pulse, ft/sec

Figure 77. Measured head velocities in sled tests with

anthropomorphic dummies and cadavers.

of occupant restraint were used and are so indicated on each

curve.

10.5 INSTRUMENT PANEL STRUCTURE PROXIMITY

Most aircraft cockpits are, of necessity, very compact. It is

necessary, for instance, for &_pilot to be able to reach var=

ious controls on the instrument panel by leaning forward no

more than 1 in. (the extent of unlocked inertia reel extension). Consequently, instrument panels must be close enough

to be reached and seen easily. Unfortunately, this usually

requires that the instrument panel and its supporting structure be placed directly above the pilot's lower legs as they

rest normally on the rudder pedals. When a seated pilot is

exposed to -G. (eyeballs-out) accelerations in a crash, the

212
lower limbs are abruptly extended longitudinally with some up- ward velocity. In this process, the lower leg usua'ly impacts

on the lower edge of the instrument panel. iepending on the

particular aircraft configuration, this contact can take place

from the kneecap down to the ankle. In -view of the high velocities associated with such flailing, disabling lower leg

injuries are common in accidents where high -G forces are

present. It is essential that desigyiers considgr using suitable energy-absorbing padding materials, frangible breakaway

panels, or ductile panel materials for structure within the

lower leg strike envelope.

10.6 RUDDER PEDAL CONFIGURATION

In certain types of aircraft accidents, the pilot's feet remain

on the rudder pedals instead of flailing upward and outward.

If the rudder pedal is a simple, bar type of arrangement, the

heel may be forced under the pedal. When the body is exposed

to a combination of vertical (Gz eyeballs-down) and longitudinal (-G eyeballs-out) forces, pelvic rotation around the

lap belt *ill almost invariably occur unless a lap belt tiedown strap is used. This pelvic rotation, which forces the

feet hard against the rudder pedals, can occur even though the

lap belt is drawn up tightly. A loose or slack lap belt aggravates the tendency toward pelvic rotation. If the forces are

great enough, a badly injured or trapped foot can result.

Therefore, it is desirable to design the rudder pedals and surrounding structure to prevent this from occurring. This is

usually done by providing a pedal capable of supporting both

the ball of the foot and the heel, and by providing a surrounding structure of sufficient strength to prevent crushing and

trapping of the lower limbs. The geometry required by MIL-STD1290(AV) (Reference 1) to prevent entrapment of feet is illustrated in Figure 78.

10.7 CONTROL COLUMNS

Control columns located in front of flight crew stations can

present a serious hazard to crewmembers if they fail- at a-ny

appreciable distance above the-aircraft floor. Such a failure

often leaves a torn, jagged stump that can inflict serious injury to a crewman should he be thrown against it during impact,

move into it as an energy-absorbing seat strokes, or come in

contact with it during egress after impact. It is recommended

that control columns be designed so that fracture due to the

occupant's striking the column will occur at a point no more

than 4 in. above the pivot point. The failure should occur in

the form of a clean break, leaving no jagged or torn edges.

Control columns that pass longitudinally through the instrument

panel are not recommended since these tend to impale the crewmembers in severe longitudinal impacts.

"213
A K -CK

Dimensions A, B, and C must be either

less than 2 in. or more than 6 in.

Figure 78. Antitorque, or rudder, pedal geometry

to prevent entrapment of feet.

10.8 SIGHTING AND VISIONIC SYSTEMS

Delethalization of the copilot/gunner (CPG) station of an attack or scout helicopter equipped with a weapon sighting optical relay tube (ORT) can present a difficult design problem.

The copilot/g-unner crewstation activities demand that the CPG

will be either in contact with the ORT eyepiece during hazardous nap-of-the-earth (NOE) flight or close to the eyepiece when

sitting in the full upright, erect position. Operational location of the CPG head, when not looking in the ORT, may be as

little as 8.5 in. from the eyepiece. Therefore, it can be expected that the CPG, when restrained by a MIL-S-58095(AV) restraint system, will contact the ORT eyepiece under nearly all

impacts over 4 G (see Figure 79). Any deformation of the bulkhead which would cause the ORT to move rearward will only further ensure head contact. Forward motion of the upper torso

after head contact with the ORT could cause spinal injury.

214

IIL,
ON1

-0

"a.14"

hi4

0

04

"A4

2154
Under NOE conditions with the CPG looking through the ORT, it

can be expected that no warning of impending impact will occur.

Regardless, any courses of action taken by the CPG to hold himself erect will probably not help in keeping his head from

striking the ORT due to head flailing and body stretch. Another factor that further decreases the distance between the

head and the ORT eyepiece is the travel of the seat as it

* strokes under crash loads.

Possible ORT hazards to the lower extremities and the torso

consist of the sharp unyielding lower structure of the ORT. In

addition, the rudder pedals may be located adjacent to the ORT.

During a crash, the potential displacement of the ORT may cause

the CPG's legs to become entrapped. A summary of typical ORT

crash hazards is presented in Table 13.

TABLE 13. POTENTIAL OPTICAL RELAY

TUBE CRASH HAZARDS

Location Type of

Hazard of injury Injury Cause

1 Head Laceration, Head strikes ORT due to

Fracture, flailing forward and downConcussion ward on impact

2 Head/Chest Crushing, Head/chest strikes ORT due

Avulsion, to ORT displacing rearward

Fracture

3 Head/Chest Laceration, CPG seat displaces downward

Crushing, and forward during energyFracture absorbing stroke. Contact

of the head/chest with

sharp edges of franged ORT.

4 Arm Laceration, CPG arms flail forward on

Fracture longitudinal impact - _

5 Lower torso Avulsion, ORT di-splaces rearward on

Laceration, longitudinal impact p ~Crushing,

S~Fracture

6 Leg Laceration, CPG leg flails forward on

Fracture longitudinal impact

7 Leg Crushing CPG leg trapped between aircraft structure and displacing ORT

216 LI/7
The cockpit should be designed to minimize the probability of

the CPG head/neck striking the ORT and minimize injury if the

CPG should strike the ORT, for both the "head-up" and "headdown" CPG positions. Some of the options available to the designer given this task are:

"* ORT Eyepiece Relocation - Consideration should be

given to reducing occupant strike hazards by moving

the ORT further away from the CPG.

"* Restraint System - The restraint system of Figure 44 4

would offer improved upper torso restraint, particularly when combined with the power-haulback inertia

reel.

"* Inflatable Restraint - Consideration should be given

to an inflatable restraint system (see Section

7.2.4). This type of restraint harness can prevent

injury to the CPG in both the erect and head-down

position by reducing slack and increasing the surface area of the body over which the harness reacts.

"* Frangible/Breakaway Features - ORT or ORT components

designed to be frangible should break away at a total

force not to exceed 500 lb. For the frangible ORT,

this force should be applied along any direction of

loading within the plane normal to the axis of the

ORT, as well as along the axis of the ORT. Breakaway point(s) of the ORT should be outside the head

strike envelope.

"* Collapsible Features - If the ORT is designed to

collapse in order to avoid injuring the CPG, the

collapse load along the axis of the ORT should not

exceed 500 lb. Figure 80 illustrates one crushable

sight eyepiece concept (from Reference 97). Two advantages of the crushable sight eyepiece are that it

is always available and, it should function regardless of head location. A helmet crash-absorber pad

would attenuate crash loads to the helmet when available crushing is expended. - -

e Power-Haulback Inertia Reel (PHBIR) - On the basis

of Air Force testing accomplished for the development of PHBIRs, the retraction time is 0.3 to

97. Fox, R., Kawa, M., and Sharp, E., DESIGNING CRASHWORTHI- NESS INTO THE YAH-63, paper presented at the Aircraft

Crashworthiness Symposium, University of Cincinnati, Cincinnati, Ohio, October 1975.

217
Normal operation

SPH-4 helmet

,•- reference

Helmet crashenergy absorber / --Pressure support

area approximately

Crash energy- 6 in.2

absorbing material

Pre loaded

prInner contours Opiaslip rings-- ofl• fa-ce mas

Optical relay Crush stroke Of face mask

tube Outer contour face mask

20-MM

eye relief

After crash stroke

Helmet crashenergy absorber

Yielded

material

Deformed tube

Figure 80. Crushable eyepiece concept. (From Reference 97)

0.4 sec, which is too slow for effectiveness in most

crashes. If this time were reduced, the retraction

velocity of the torso would have to be increased

considerably over the current limit.of 9 ft/sec. A

retraction velocity greater than this is not recommended due to the lack of human tolerance data on

this type of loading. In a crash with a single pulse

218

• m " m ' n ' " m m m ° ' • ' " "I'
of say 30-G peak and 50-ft/sec velocity change, the

retraction velocity should be approximately 25 ft/sec;

therefore, thle known tolerance limits would be exceeded at the higher velocity. In summary, the PHBIR,

as currently qualified under both Air Force and Navy

military specifications, requires excessive time to

position the tcrso by crash sensing. To be fully

effective, the system should move the torso into position in approximately 0.06 sec, but the resulting

acceleration would exceed known human tolerance

limits. The primary crashworthiness advantage of the

PHBIR would be as a manually activated tightening device for the head-up CPG position; the PHBIR offers

only limited advantage for the head-down CPG position.

10.9 ENERGY-ABSORBING REQUIREMENTS

FOR COCKPIT AND CABIN INTERIORS

10.9.1 General

To minimize occupant injury, the acceleration experienced during secondary impacts of the occupant with surrounding structures must be reduced to a tolerable level. The areas of contact to be considered for energy absorption include instrument

panels, glare shields, other interior surfaces within the occupant's strike envelope, and seat cushions. A padding material

should not only reduce the decelerative force exerted on an impacting body segment, but should distribute the load in order

to produce a more uniform pressure of safe magnitude.

As an example of the need for an energy-absorbing system to

possess both these characteristics, consider the case of head

impact. Head injuries sustained from impact may be grouped in

two general categories. The first is skull fracture with its

inherent brain damage and danger to life. The second is injury to facial tissue and bone structure with a lesser pi:obability of brain damage.

A system that is to absorb the energy of an impacting head

should cushion the head to prevent skull fracture or penetration from protruding objects as a result of decelerative forces.

It should also distribute the forces-to minimize injury to tissue and bone Etructure. The cushioning material used must effect low peak deceleration and low average stress. Figures 81

and 82, taken from Reference 98, indicate the impact behavior

98. Lee, W. M., and Williams, B. M., CUSHIONING AND LOAD DISTRIBUTION PERFORMANCE OF PLASTIC FOAMS, Paper No. 700453,

Society of Automotive Engineers, Inc., New York, May 1970.

21

.1 * .*.-. . .r-s..
175-

Semirigid urethane

15o foam (9 lb/ft 3 )

S125

0

3 Polystyrene beadboard

a) 3 C.) (1 lb/ft )

10

E 75-

V

4)

4 50- _____ i____ ___

\,Polyethylene

foam (2 lb/ft

25 (Te

(Testing at room temperature)

0 5 105 20 25 30

Impact speed, mi/h

Figure 81. Impact behavior (headform deceleration

versus speed) of three padding materials.

of three plastic foams. The foam sample specimens used to obtain these data were 6 in. thick to minimize any bottoming-out

effect. Although the semirigid urethane appears to be a fair

cushioning material, it does not distribute the load as well

as the materials with which it is compared. A fair cushioning

material is not necessarily an effective load distributor.

Both criteria must be considered in the selection of a material

that is to provide impact protection for the head.

220

I-lrrl rl --. 1- ml-rr'11*l'
. .i 120. I I

(Testing at room temperature) _ _I _ I

S100 Semirigid urethane or

foam (9 lb/ft 3 )"

- . S80

.U 6 0 -P

40

60 - -3

0 foam (2 lb/ft

20 2

fir Polystyrene

1/' beadboard

Z (1 lb/ft)

0 5 10 15 20 25 30

Impact speed, mi/h

Figure 82. Impact behavior (average dynamic stress

versus speed) of three padding materials.

In addition to protecting bone structure and facial tissue, the

energy-absorbing system must also afford protection- against

intracranial lesions. Cerebral concussion, and the loss of

consciousness which often accompanies it, may occur if the head

is subjected to excessive decelerative forces. Mattingly, et

al. (Reference 99), in discussing possible intracranial lesions

and cerebral trauma including concussion, swelling, contusion,

laceration, and hematoma, conclude that in order to prevent

99W Mattingly, T. E., e al., INVESTIGATION OF VIBRATION AND

IMPACT PROTECTION OF THE HUMAN HEAD AND NECK, Northrop Corporate Laboratoriesi AMRL Technical Report 69-112, Air

Force Systems Command, Wright-Patterson Air Force Base, Ohio, December 1969, AD 702124.

221

I.......~..
head injury, materials must be carefully selected to absorb

and attenuate the energy of impact. The material must reduce

the level of acceleration, the rate of onset, and the amount

of energy transmitted to the head.

10.9.2 Types of Padding Materials and Properties

The most useful types of materials for energy-absorbing padding are plastic foams. A foamed plastic is usually totally unlike the same plastic in the solid: unlike in properties,

in processing, and usually in applications. Three steps are

involved in producing a cellular structure in a polymer: 1)

preparation of polymeric material into a viscous liquid state,

2) introduction of fine bubbles of gas to produce expansion,

and 3) solidification of the foamed plastic to stabilize the

foamed structure. The particular process used in manufacturing

foam materials has a direct effect on their properties and can

result in products of the same chemical composition being very

different in performance.

10.9.2.1 Material Form: The form in which the foam material

is commercially available influences its adaptability to vehicle applications. Slab and molded foams are often used in the

construction of instrument panels and seat systems. Differences in properties due to varying the form should be considered in the selection of a material. For example, Figure 83

shows the variation of minimum tensile strength versus product

density for polyethylene foam in sheet and plank forms.

10.9.2.2 Classification of Foams: Foams can be described as

flexible or rigid. A flexible foam recovers when deformed,

whereas a rigid foam cannot sustain multiple impacts. Flexible foams are most widely used in situations where energy absorption is important.

Another method of classifying foams is open-cell or closed-cell.

An open-cell foam contains individuAl cells that interconnect

with the others, while in a closed-cell foam individual cells

are completely enclosed by a wall of plastic.---

Plastic foam materials also can be classified according to

their chemical composition. Several energy-absorbing plastic

foams and some of their typical applications are listed in

Table 14.

10.9.2.3 Material Properties: The selection of a foam material for vehicle energy-absorbing applications involves an evaluation of its processabilityl its mechanical, thermal, and

222
*14 heetPlank • • ~SheetPln

No100 L

op

4 80

w /

Prodc d

Figure 83. Minimum tensile strength versus product

density for polyethylene sheet and plank. (From Reference 100)

chemical propertiesi as well as its cost. Along with the primary foam materials, the characteristics of adhesives and surface coatings must be considered, particularly with respect to

emission of smoke and toxic vapors. The characteristics of

suitable materials for such use are listed below:

e Adaptability and ease of processing

* High energy dissipation

a Effective load distribution

* Low rebound

o Temperature insensitivity

e Low w0ter absorption

200. Fradosj. j., PLASTICS ENGINEERING HANDBOOK OF THE SOCIETY

OF THEt PLASTICS INDUSTRY, I*NC., Van Nostrand Reinhold Co., New York, 1976, pp. 499-567.

S~223.

-. 4a). -

'II

-'U m -..
e Resistance to chemicals, oil, ultraviolet radiation,

and sunlight

* Nontoxic, fume generation

* Favorable flammability rating

* Minimal smoke generation

* Durability and long life

* Cost competitive I

* Aesthetic

TABLE 14. ENERGY-ABSORBING PLASTIC FOAMS

AND SOME TYPICAL APPLICATIONS

r. Semirigid and flexible urethane foam

Aircraft, automobile, and furniture seat cushions,

safety padding, arm rests, sun visors, horn buttons, bedding, carpet underlay, packaging delicate

products.

2. Polyvinylchloride foam

Crash padding in automobile head liners and sun

visors, flooring, shoe soles and heels, automobile door panels, seating upholstery sealants,

gaskets, bumperstock.

3. Polystyrene foam I

I

Insulation, packaging.

4. Expanded rubber

Bus and subway seat cushions, truck and ship

mattresses, gaskets, hose insulation.

5. Polyester foam

Short-run, custom-type seat cushioning.

6. Polyolefin foam

Packaging, gasketing, water sports equipment, rug

underlay, athletic padding, antivibration padding.

Foam materials are most often characterized by the mechanical

properties listed below, where it may be noted that several

of the properties apply only to rigid or flexible foams. For

224
example, compressive strength is not relevant in considering flexible foams. The SA compression-set test, on the other hand,

applies only to flexible materials.

• Density

* Tensile strength

* Tensile modulus

* Compressive strength

• Compressive modulus

* Flexural strength

* Flexural modulus

* Tear strength

* Compression set

* Compression deflection

e Elongation

* Rebound

e Hardness

* Impact

Properties of possible interest in selection of a material for

energy-absorbing applications are presented in Table 15 for

several applicable materials (data taken from References 100

through 103).

101. ENSOLITE, Publication ASP 9997, Expanded Products Depart--

ment, Uniroyal, Inc., Mishawaka, Indiana, 1977.

102. TEMPER FOAM, Form TF-20, Edmont-Wilson, Division of Becton, Dickenson and Co., Coshocton, Ohio, 1975.

103. Sundquist, D. J., POLYOLEFIN FOAMS, Monographs on Plastics, Vol. 1, Part 1, 1972, pp. 193-295.

1 225

k

* ..- _ ." . . .-- - ;.. _..*, ."....- • .-

Sllim
ad to o Ad.

40 L I '4~'u * v- .in

U'-403 4 j 3

0.0

.4 0~o I

0U~ in V

0 in 0 N nP14 (hN n

00

>44). 4

U.. 0jN

o 0 H 41 0 n0 . 0

4, -,4 .,4 4j. 0

V. r. 4 0.4 41

a04

0f cc to4 4m in 0 co5

* 4. 4 ~ . 4 J ~ j.22 61
10.9.3 Standard Test Methods

ASTM standard test procedures. are widely used by manufacturers

to specify various properties of a particular type of material.

Table 16 summarizes ASTM test methods and specifications for

flexible cellular plastics that provide a basis for comparison

of materials. Here it may be noted that most ASTM tests involve simple tests, whereas the operational environment involves dynamic loading and more complex conditions.

In particular, ASTM D 1564-71 describes "Standard Methods of

Testing Flexible Cellular Materials-Slab Urethane Foam" (Reference 104). Among other tests, there are compression-set and

load-deflection tests. In the compression-set test, the method consists of deflecting the foam specimen under specified

conditions of time and temperature and noting the reduction of

specimen thickness after removal of the load. The compression

device consists of two flat plates larger than the specimen.

In the load deflection test, one method consists of measuring

the Indentation Load Deflection (ILD) value, which is the load

necessary to produce a specified 25-perceft or 65-percent indentation in the specimen under a 50-in. circular indenter

foot. Acceptable deflection rates range from 1.0 to 15.0

in./min. A second method, which uses the same indenter, obtains the deflections under specified loads of 4.45, 111, and

222 N (1, 25, and 50 lb) during loading and 111 N during unloading. Theae deflections are reported as Indentation Residual Gage Load (IRGL) values. The latter method, which involves

indentation to specified loads, is intended for use with seat

cushion materials.

The above tests provide results that specify the material, but

do not necessarily portray its performance under actual impact

situations. A simple dynamic drop test, such as ASTM D1596-64

(1976), "Standard Test Method for Shock-Absorbing Characteristics of Package Cushioning Materials" (Reference 104), more

closely simulates actual impact conditions. An acceleratiortime curve is obtained by mounting a transducer on the dropping

head. The parameters evaluated are peak deceleration and the

dynamic set of the specimen. This method allows the test

parameters to vary and yet is simple enough to ensure repeatability among different test facilities. In a drop test, the

test parameters are: the drop height that determines the impact velocity, the weight and surface area of the impactor,

and foam thickness.

104. Lukens, R. P., et al., 1977 ANNUAL BOOK OF ASTM STANDARDS,

American Society for Testing and Materials, Easton, Maryland, 1977, Parts 20, 38, 48.

227

- 4 -... . - . . . . . .. • _
r

TABLE 16. SUMMARY OF ASTM TEST METHODS AND

SPECIFICATIONS FOR FLEXIBLE

CELLULAR PLASTICS (REFERENCES

104 AND 105)

D1564-71* Testing Flexible Cellular MaterialsSlab Urethane Foam

D1667-76* Specification for Flexible Cellular

Materials - Vinyl Chloride Polymers

and Copolymers (Closed-Cell Sponge)

D1565-76* Specification for Flexible Cellular

Materials - Vinyl Chloride Polymers

and Copolymers (Open-Cell Foam)

D1055-69* Specification for Flexible Cellular

(1975) Materials - Latex Foam

DI056-73* Specification for Flexible Cel.ular

Materials - Sponge or Expanded Rubber

D3575-77 Testing Flexible Cellular Materials

Made From Olefin Plastics

D1596-64* Test for Shock-Absorbing Characteristics

(1976) of Package Cushioning Materials

D2221-68* Test for Creep Properties of Package

(1973) Cushioning Materials

D1372-64* Testing Package Cushioning Materials

(1976)

D696-70* Test for Coefficient of Linear Thermal

Expansion of Plastics

E143-61* Test for Shear Modulus at Room Temp-

(1972) erature

D412-75* Tests for Rubber Properties in Tension

D1433-76" Test for Rate of Burning and/or Extent

and Time of Burning of Flexible Thin

Plastic Sheeting Supported on a 45-

degree Incline

D1692-76 Test for Rate of Burning and/or Extent

and Time of Burning of Cellular Plastics Using a Speciman Supported by a

Horizontal Screen

*Indicates that the standard has been approved as

American National Standard by the American National

Standards Institute.

228
Other standard test procedures include SAE J815, "Load Deflection Testing of Urethane Foams for Automotive Seating," as described in Reference 106. This test points out the factors of

interest in testing materials for vehicle seat cushions: the

thickness of the padding under the average passenger load, a

measurement that indicates the initial softness, and a measurement that indicates resiliency. SAE J815 determines load versus deflection by measuring the thickness of the padding under

fixed loads-of 1 lb, 25 lb, and 50 lb with a circular indentor

foot (see Reference 105).

Also, SAE J388, "Dynamic Flex Fatigue Test for Slab Urethane

Foam" (Reference 107), describes procedures for evaluating the

loss of thickness and the amount of structural breakdown of

slab urethane foam seating materials. A test specimen is measured for thickness under a specified load and, subsequently,

subjected simultaneously to compressive and shear deformation

under a constant load for a specified number of cycles. In

the constant load height measuring test, a flat, circular

indenter foot of 50 in. wilh loads from 1.0 to 75.0 lb is deflected at rates from 2 to 8 in./min. The constant load dynamic fatigue apparatus uses rollers in a more complicated

setup.

SAE J921, "Motor Vehicle Instrument Panel Laboratory Impact Test Procedure-Head Area," describes a test procedure for evaluating the head impact characteristics of such areas as instrument panels (Reference 96). An SAE J984 headform with

an effective weight of 15.1 lb is impacted at specified positions. The parameters evaluated are the impact velocity, the

acceleration-time history of Lhe htuoiorm, and the start of

impact, with optional measure,4Ient of the rebound velocity and

the headform dynamic displacement.

105. ENCYCLOPEDIA OF POLYMER SCIENCE AND TBCHNOLOGY - PLASTICS,

RESINS, RUBBER, FIBERS, John Wiley and Sons, Inc..,- New

-York, 1965, Vol. 3, pp. 98-126.

106. SAE Recommended Practice, SAE J815, LOAD DEFLECTION TESTING OF URETHANE FOAMS FOR AUTOMOTIVE SEATING, SAE Handbook

1979, Part 2, Society of Automotive Engineers, Inc., Warrendale, Pennsylvania, 1979, p. 34.31.

107. SAE Recommended Practice, SAE J388, DYNAMIC FLEX FATIGUE

TEST FOR SLAB POLYURETHANE FOAM, SAE Handbook, 1979,

Part 2, Society of Automotive Engineers, Inc., Warrendale,

Pennsylvania, 1979, pp. 34.28-34.30.

229

-- A, ---- -------.. -.. ~.. . . . . .
S10.9.4 Research on Materials for Energy-Absobing plications

Static tests that deviate from ASTM test procedures and simple

dynamic tests that are intended to grossly simulate crash conditions have been performed by manufacturers and users with

different types of materials. Several of their approaches and

their energy-absorption criteria are discussed below.

10.9.4.1 Acceptable Stress-Strain Characteristics: Haley, et

al., have investigated design criteria for pacding materials,

as described in Reference 94. According to their conclusions,

energy-ablorbing materials with stress values between 40 end

80 lb/in. at 50 percent strain would offer reasonable survival

potential for head impacts on flat surfaces at velocities of up

to 20 ft/sec with a padding thickness of 1.5 in. More recent

unpublished data gathered by the Army's Aeromedical Research

Laboratory (USAARL), Fort Rucker, Alabama, indicates that the

above stress values are too high because the values were based

on the compressive strength alone while it is probable that

tensile stresses and shear stresses around the periphery of the

compressed areas play a large role in the total force resisting compression. Regardless of the stress distribution in the padding material, the USAARL research has sholn that the stress

level should fall between 30 and 4521b/in. for padding less

than I in. thick and 20 to 30 lb/in. for padding greater than

I in. thick in order to prevent peak G pulses from exceeding

the tolerance values stated in Volume 11. These crush strength values, as illustrated in Figure 84, are recommended. These

levels are expected to prevent unconsciousness (within the

limits of the crush depth) for head impacts. The lower stress

level for the thicker padding is based on: (1) the average design decelerative level must be reduced as the depth of the

padding and concomitant time duration are incr'.eased to meet the

known tolerance limits stated in Volume II, and (2) a larger

area of foam is crushed as the head sinks into the thicker pad.

Use of a padding as proposed in Figure 84, is intended to limit

head peak G values tc 160 for the thin pads and 120 G for the

thicker padding.

The criteria of Figure 84 are to be satisfied by the padding

material over the entire anticipated operating temperature

range if the potential for survival is to be maintained. Practical considerations and risk analysis, however, may reduce the temperature range requirements. Figure 85, taken from

Reference 108, indicates the temperature dependency of the

stress-strain properties of a particular foam material. It

illustrates the variation experienced by many padding materials

108. PACKAGING WITH ETHAFOAM, Publication No. 172-221-IOM-767,

Dow Chemical Company, Midland, Michigan, revised 1966.

230
. 40Ot~l • '

0 

~ vb.:cept~hi.

CUAhGceptAbl. Acceptebim

20

D) NrO paddinq thicknes1 lies thn on. Lnch For padding thickmi grq ater thah one inch

_ _ _ _ _ _ _ _ I ,- I ,

Str,!n, petrent

Figure 84. Recommended stress-strain properties for padding

material for head contact, with cushion thickness

of at least 1.5 in. (From Reference 94)

and indicates that temperature sensitivity must be considered

as a padding material selection criterion.

Stress-strain curves for a polyurethane-foamed plastic are

shown 4n Figure 86. The curves are taken from Reference 109

and indicate that a 6-in. tnickness of the foam with a density

of 5 lb/ft will satisfy the criteria of Figure 84 (superimposed as a crosshatched area) over at least part of the oper- ational temperature range. The lowest impact velocity used to

obtain the data of Figure 86 was 50 ft/sec. A weight of 295 lb

impacting at this velocity requires the absorption of over

11,000 ft-lb of energy by the padding material. This requirement is obviously considerably more demanding than that of 90

ft-lb of energy at an impact velocity of 20 ft/sec, as described above. Further work with variations of this foam may

yield a material that will satisfy the design criteria with a I

thickness of 2 in. or less.

10.9.4.2 Bioengineering Approach to Material Evaluation:

Daniel investigated the injury-reducing functions of crash padding, considering strength of skull segments, as described in

Reference 110. He concluded that because the cranial vault

109. ENGINEERING DESIGN HANDBOOK, DESIGN FOR AIR TRANSPORT

AND AIRDROP OF MATERIAL, AMC Pamphlet No. 706-130, U. S.

Army Materiel Command, Wahhington, D. C., December 1967,

AD 830262.

110. Daniel, R. P., A BIO-ENGINEERING APPROACH TO CRASH PADDING, Paper No. 680001, Society of Automotive Engineers, Inc.,

New York, 1968.

231

________.._________ .

m m _______ - - ... ...
-A

20

18 ._

16.

14 - ,

S12 ____

1i0 __ -40OF _

+73 F

~' 8

6

2

0-

0 10 20 30 40 50

Strain, percent

Figure 85. Stress-strain curves for polyethylene foam. (From Reference 108)

(above the eyebrows) is strong under localized impact, padding used for protection of this region has the primary-function of

energy absorption to reduce the possibility of brain damage.

On the other hand, padding for facial protection should distribute the impact load over the weaker facial bones, and required energy absorption would be provided by the supporting structure. His suggested evaluation criterion for energy- absorbing materials, based on t program of 91 impact tests, is

illustrated in Figure 87. For any given material, plotting on

these curves the results of a test conducted according to the

given parameters would enable the determination of a material

"efficiency," where a 100-percent efficiency would correspond to the deceleration achieved by an ideal square-wave energy

232
M0 I

so 0

70 - all P": 00

,//7

70 -

-.C 4 k lI 41

N, 50

Wn 40

20

no Claim* I / 0en~tytong Drolped Velocity I (rad) Size

(Ib/Itl)1 Tcqts Ye.ar 1.)I) ' tsec) i n.)

40 4 1S ;95 I 51, 74x2lxi

10 3- 0 1 I f ) N/A Qus . 1 7

x1,sxO.4,

2.0 0. 2 1979 N/A - ,,e- 2.0Na.sox.O

*Foem taken from foa.-itn-place helmot.

rFo m apec imen produced from U.S. Air Force Helmet Foaemtiq Kit

witI. equal tartm of Prlyol an-I Iaocyanate.

0 I , I I 1

0 20 40 60 80 100

Compressive stra!.n, percent

Figure 86. Effect of density on stress-strain curves

for polyurethane-foamed plastic.

233

10 o

- , : . : ..],1• ;-4
4J 4

0 4

0I 'P ,4400 

Q -r4 U 0 t$4n 0

~4 04)- iv ~ N ~

'04 41qb .0

to x~JUU CH

4' )

U 00

* 0C

- "4

Ln I 02 

en~ N

'UOTvJ9aO- waope

234 '4.
absorber of the given thickness. According to Daniels, energyabsorbing materials might be selected on the basis of maximum

efficiency.

Evaluation criteria for load-distributing applications, which

are illustrated in Figure 88, are based on the following assumptions:

* A load-distributing pad should permit the face to

penetrate its surface relatively easily and then

maintain a cushioning layer of foam between the base

and the underlying structure during collapse of the

understructureI

* The understructure should deform at close to the 80-G

(1200 lb) face tolerance level expressed in both SAE

J885 and Federal Motor Vehicle Safety Standard 201

(References 111 and 112, respectively).

10.9.4.3 Energy-Absorbing Efficiency Calculations.- Th- energyabsorbing characteristics of foamed polymers were mathematically calculated by Rusch from low-speed experimental data for

compressive strain and modulus (Reference 113). Materials were

characterized by three parameterst energy-absorbing efficiency,

impact energy per unit volume divided by foam modulus, and the

maximum decelerating force per unit area divided by foam mod- ulus.

An ideal energy absorber would provide a constant deceleration

from an initial speed, vii for 100 percent of its thickness, h

The maximum deceleration for an ideal absorber is then given by

dm vi /2h d~i (36)

111. SAE Information Report, SAE J885, HUMAN TOLERANCE TO IMPACT CONDITIONS AS RELATED TO MOTOR VEHICLE DESIGN, SAE - Handbook 1979, Part 2, Society of Automotive Engineers,

Inc., Warrendale, Pennsylvania, 1979, pp. 34,114-34.117.

112. U. B. Code of Federal Regulations, Title 49, Part 571:

MOTOR VEHICLE SAFETY STANDARDS, 201, OCCUPANT PROTECTION

AND INTERIOR IMPACT, Government Printing Office, Washington, D. C., (Rev.) 1978.

113. Rusch, K. C., IMPACT ENERGY ABSORPTION BY FOAMED POLYMERS, Journal of Cellular Plastics, Vol. 7, No. 2, 1971,

pp. 78-83.

235

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The energy-absorbing efficiency, K, is defined as the inverse

ratio of the maximum deceleration exhibited by a real material,

dm, to that for an ideal material of equivalent thickness, dmi,

12

K vi i /2hd m ((37)

Generally, K is expressed as a function of the impact velocity.

At low v , the impact energy is small relative to the stiffness

-of the foam, the degree of penetration is small, and K is low.

r. At high vi, the impact energy is large relative to the stiff-

[ ness, the impacting body "bottoms" on the understructure, and

K is low. At some intermediate vi, K exhibits a maximum.

The optimum material is one for which: (1) the K versus v.

curve is as broad as possible, (2) Kmax is close to unity, and (3) K occurs at the most probable vi for the particular applicati1

On the basis of his calculations, Rusch stated the following conclusions: (1) the energy-absorbing characteristics of a

brittle foam are superior to those of a ductile foam; (2) the

optimum energy-absorbing foam has a large cell size, a narrow

cell size distribution, and a minimum number of reinforcing

membranes between the cells; and (3) foam composites offer no

significant advantage over a single foam.

10.9.4.4 Composite Foam System: Brooks and Rey (Reference 114)

found that a composite could be formed combining the high energy dissipation of polystyrene beaded foam with the load-distributing effects of semi-rigid urethane. Simple dynamic tests

consisted of dropping a 6-1/2-in. diameter aluminum hemispherical headform weighing 15 lb at impact velocities up to 30 mi/h

(44 ft/sec). As shown in Figure 89a, the urethane exhibits the

lowest level of headform acceleration during impact. On the

other hand, the polystyrene exhibits the lowest level of penetration as shown in Figure 89b. The urethane can be said to

absorb the least amount of energy, as indicated by the highest

rebound value in Figure 89c. - -

In-small-scale static tests, 2-in. cubes were compressed to

70-percent deflection and then relaxed with an Instron testing

machine at 2.0-in./min crosshead speed. Figure 89d shows the

relative energy absorption of the three materials tested, indicating the composite foam as a compromise between polystyrene and semi-rigid urethane foam.

114. Brooks, J. D., and Rey, L. G., POLYSTYRENE-URETHANE COMPOSITE FOAM FOR CRASH PADDING APPLICATION, Limited Publication, Low Chemical of Canada, Sarnia, Ontario.

237

.. .

___-_____, -- -.- .. .. - .. . .. --- , ,.4 . .
130

Polystyrene 7 Z

110 Urethane

Compos.te

90

00

70

30 o

35 - 0 2 3

i5 20 25 30

Velocity, mi/h i 20 25 30

(a) Head form acceleration Velocity mi/h

(b) Headform penetration

40 - - 60 -- Polystyrene

----Composite --- ........ Urethane

35 4SUrethane .,/ 

030 2 30 -

""'00

M 0

S25 -- 15 - 4 Composite /.

"20 // ::

- Polystyrene 0 10 30 50 70

Compression, percent

-W (d) Compression load- 15 20 25 30 deflection curves

Velocity mi/h

(c) Headform rebound

Figure 89. Response of composite foam, compared with urethane,

and r-1-styrene. (From Reference 114)

238

ii
SA 10.9.4.5 Specific Energy and Relative Energy-Absorption Ratio:

Reference 115 discusses performance parameters of Dow composite foam in energy-absorbing applications. It was concluded that,

on the typical response curve for a compression test, where the

area contained within the hysteresis loop shown in Figure 90

is directly related to the energy absorbed, three performance

parameters can be defined: the specific energy absorbed at

maximum strain, the relative energy-absorption ratio, and the

maximum stress. The total energy absorbed at maximum strain is

the sum of areas A and B. When this total energy is expressed

in terms of a unit volume (or unit weight), the quantity becomes the specific energy absorption at maximum strain. The

ratio of area A to the sum is the relative energy-absorption

ratio, which is a measure of the amount of energy actually dissipated during compression. In effect, it corrects the performance parameter for the energy that is momentarily stored.

The maximum stress is usually the stress at maximum strain.

Exceptions to this occur wthen some rigid cellular materials are

compressed and a spike is observed during the initial stage of

compression. Maximum stress levels are directly related to the

deceleration that the impacting object sustains.

ra ý4) I o ~

L~n B

Deflection

Figure 90. Typical response of plastic

foam to compression test.

115. PERFORMANCE PARAMETERS OF DOW COMPOSITE FOAM IN ENERGY

ABSORBING APPLICATIONS, Dow Chemical Company, Midland,

Michigan.

239
Melvin and Roberts (Reference 116) measured the specific energy

absorbed and the relative energy-absorption ratio for the materials listed in Table 17 using three speeds: 20, 2000, and

13,000 in./min. Their results are summarized in Table 18 and

Figure 91, from which they concluded that the majority of foams

do not exhibit marked increases in properties with increasing

test speed. The vinyl foams, which exhibit dramatic increases,

are the exceptions.

TABLE 17. MATERIAL SUMMARY (FROM REFERENCE 116)

Initial strain rate

Density Specimen (sec )

Material and 3 height

code number (lb/ft3) (in.) Speed 1 Speed 2 Speed 3

Polyethylene E-1 2.34 2 0.17 17 100

Polyethylene E-2 6.65 1.5 0.22 22 150

Polyethylene E-3 9.05 2 0.17 17 100

Polystyrene S-1 1.09 2 0.17 17 100

Polystyrene S-2 3.35 2 0.17 17 100

Polystyrene S-3 1.21 2 0.17 17 100

(pelletized)

Polyurethane U-1 1.53 2 0.17 17 100

(rigid)

Vinyl V-1 7.35 1 0.33 33 220

Vinyl V-2 7.25 1 0.33 33 220

Vinyl V-3 5.04 1 0.33 33 220

Cork C-l 11.5 1.5 0.22 22 150

116. Melvin, J. W., and Roberts, V. L., COMPRESSION OF CELLULAR

PLASTICS AT HIGH STRAIN RATES, Journal of Cellular Plastics, March/April 1971, pp. 97-100.

240 "
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2 10 0 5-10 0 0 0000 0000 00

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Crsha vlct, n/i

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strain versus crosshead velocity.

10.9.4.6 1namic Property Index: Fan (Reference 1171 developed techniques for simulating the force-penetration properties

of viscoelastic materials based on results of pendulum impact

tests on polyurethane foam. The dynamic force-penetration relationship of polyurethane can be approximated by a function of

three variables: penetration-thickness ratio, sample thickness,

and impact velocity. Fan suggested a criterion for energy absorption expressed as the dynamic-property index:

I E dd/Gm (38)

117. Fan, W. R. S., A SIMULATION OF THE DYNAMIC PROPERTIES OF

ENERGY-ABSORBING MATERIALS, 1970 International Automobile

Safety Conference Compendium, Society of Automotive Engineers, Inc., New York, 1970, pp. 1075-1083.

242

I . ..... ... ... ..... .
where I - dynamic-property index of the material.

Ed = the amount of energy dissipation by the foam material during impact.

Gm Gm a the maximum deceleration measured at the impactor during impact.

A high index value implies a high degree of effectiveness. The

dynamic-property index of a material varies with the test conditions. The material rated as the most effective in a certain case is not necessarily the most effective material in

other cases.

10.9.4.7 Dynamic Crushing Pressure: Furio and Gilbert (Reference 118) c~nducted a series of drop tests with low density (2 lb/ft ) urethane foam using a flat impactor weighing 729 lb at a drop height of 45 ft.

The dynamic crushing pressure, Pcr, which is the product of

the weight of the impact mass and the acceleration divided by the impact area, is shown in Figure 92 as a function of temperature and velocity for two samples of identical dimensions.

The increase in pressure with velocity is attributed to the

fact that the entrapped gas must escape in order for the foam

to collapse. Under dynamic loading, the gas cannot escape fast

enough, and a higher pressure results.

10.9.5 Application of Padding Material

In the absence cf data for extremity impacts, it is assumed

that padding material that is suitable for head impact protec- tion will be suitable also for protecting extremities. Extremity impacts are not likely to have the potentially severe effects of head impacts. It is suggested that areas within the

extremity strike envelope having radii of 2 in. or less be pad- ded and that such padding have a minimum thickness of 0.75 in.

Caution must be exercised in padding sharp edges and corners.

Padding installed in a manner that allows it to-be broken away

from the corner or cut through by sharp edges offers no protection. It is recommended that edges and corners to be padded have a minimum radius of 0.5 in. prior to padding. A definite

volume of the padding must be crushed to absorb the initial

kinetic energy of the head and protective helmet.

118. Furio, A. J., Jr., and Gilbert, W. E., IMPACT TESTS OF

URETHANE FOAM, Report No. NSRDC 4254, Naval Ship Research

and Development Center, Bethesda, Maryland, January 1974,

AD 775903.

243

r
S60

24 x 24 x 12 in.

Ambient

24 x 24 x 12 in. I 140 4// Cold (-100F)

• • ~30...

S 20....-

UI I

>11 10

0• 0

10 20 30 40

Impact velccity, ft/sec

Figure 92. Dynamic crushing pressure versus

impact velocity for tests at two

temperatures.

10.9.6 Ductile Materials

In cases where the use of padding material is impractical or

the thickness allowed is inadequate to provide the necessary protection, ductile energy-absorbing materials or frangible breakaway panels should be used where possible. Window and door frames, control columns, electrical junction boxes, etc.,• should be designed with large radii (I in. or more) rather than

with sharp edges and corners. -

Swearingen concluded in Reference 119 that at impact velocities

of 30 ft/sec against rigid structure padded with materials even

6-in. thick, unconsciousness, concussion, and/or fatal head

injuries will be produced. Where possible, a combination of

119. Swearingen, J. J., EVALUATIONS OF VARIOUS PADDING MATERIALS FOR CRASH PROTECTION, FAA Technical Report AM. 66-40,

Federal Aviation Administration, Civil Aeromedical Institute, Washington, D. C., December 1966, AD 647048.

244

4-!
deformable structure and padding material should be considered

to absorb the impact energy and to adequately distribute the

forces over the face. Surfaces to which this combination

shuuld be applied are instrument panels, seat backs, bulkheads,

and any other structure the head may impact during the crash

sequence.

A

I

Ii

245 1.

.. . - . - .- . ... .
REFERENCES

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246

I.
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i IIIIII - I 'III ' -• - -* r - " i--
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* I . ... . - -,- - - - -• .* .- 
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) 251

* 4 . -: . .. . .. . . .. .. . . .. .
I

4

REFERENCES (Continued)

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2

............................... *. ,V ..
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256

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REFERENCES (Continued)

103. Sundquist, D. J., POLYOLEFIN FOAMS, Monographs on Plas- I tics, Vol. 1, Part 1, 1972, pp. 193-295.

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2.5

258

S . A .
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260

, -•• _________-_________...__________..._i• •,'• -
BIBLIOGRAPHY (Continued)

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261

t~.!

- -.- -.-. •
BIBLIOGRAPHY (Continued)

Robbins, D. H., et al., DEVELOPMENT AND TESTING OF INTEGRATED

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262
BIBLIOGRAPHY (Continued)

'S

Underhill, B., and McCullough, B., AN ENERGY-ABSOPBING SEAT

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AD A038762.

t

263

I UC

A .. - .i
INDEX

Acceleration, definition 26

Adjustment, seat 88

Adjustment hardware, location 146, 163

Aftward load, definition 29

Aftward strength 172

Air Force head/spine model 102

Aircraft coordinate system 25

Airframe, energy absorption by 40

Anchorage, restraint system 67, 141, 144

Anthropomorphic dummy

definition 31

seat testing 188

Attitude

definition 25

landing gear 43

Back tangent line

definition 53

limit for comfort 37

tolerance 172

Body block, static testing 185

Bolted connections

restraint systems 147

seats 50

Bottoming, definition 36

Buckle

aircrew 130, 163

single-point release 139

troop/passenger 134, 163

Bulkhead

definition 36

distortion 46, 58

warping for test 182

Buttock reference line 34, 141

Buttock reference point 35

Cable energy absorber 121

Castings 50,-54

Ceiling-mounted seat 45, 51

Clearance for-seat stroke 42

Collapse

definition 36

prevention of 64

Combined load, definition 29

Comfort 37

cushion 125

restraint system 137

seat adjustment 38

vibration 39

264

,•[ii II~iii I il II!11 li-i1-1111- II lil -- -- l l
INDEX (Continued)

Computer simulation 93

Air Force head/spine model 102

CVS 99

one-dimensional 103

PROMETHEUS 100

SOM-LA 95

Controls 213

Crash environment 40, 43

Crew station geometry 33, 37, 39

Cushion

comfort 38, 125

desirable features 126

dynamic overshoot 127

energy-absorbing 127

filled 127

net-type 128

ventilation 38, 125

CVS computer program 99

Damping, comfort influenced by 37

Deceleration, characteristic waveform 85

Design conditions for seats 44

Design eye position

definition 33

seat adjustment 39

Downward load

definition 29

strength requirement 172

Ductility, seat materials 49

Dummy

definition 31

seat testing 188

Dynamic overshoot

cushion 127

definition 29

restraint system 69, 187

typical seat response 89

Dynamic response

empirical 85

theoretical 82

Dynamic testing

data acquisition 193

litters 201

seats 18

Effective weight

calculation 175

definition 32

limit load 82, 173

seat design 166, 173

265

- ~---

,*, -
INDEX (Continued)

Efficiency, Emergency release, energy absorber restraint 81, 235 system 139

Energy absorber

cable 121

definition 36 design loads, crewseats 173, 175 design loads, troop seats 178 desirable features 106 honeycomb 118 inversion tube 112

notched 110 optimization of waveform 91 rate sensitivity 175 rod Pull-through tube 121 rolling torus 114 tension pulley 123

tube tension flaring tube 119 121 variable-load 92, 177

S~Energy-absorbing wire bending 109stroke

Energy-absorbing lateral or longitudinal 42 webbing 159 Energy absorption airframe contribution 40 landing gear 40, 174 optical relay tube 217 plastic foams 219, 230 restraint webbing 159 seat contribution 41, 69, 174 seat stroke example 70

Equipment stopping weights, distance 67 seat design 167 Facial injury, prevention 232 Failure, definition 36

Floor

Attachment 45, 51, 64 distortion 46, 52 warping for test 54, 183 Forward load

definition 28 strength requirements 168 Frangible tube energy absorber 121 Frequencies, critical 39 Friction, in seat stroke 173 Head impact hazards 210 Headrest cushion 128 Heel rest line, definition 35

266

L J L > 

- _ _ _ _ _ _ _ 

_ _ 

_
INDEX (Continued)

Honeycomb energy absorber 118

Horizontal vision line, definition 33

Human body coordinates 30

Human tolerance

definition 31

head impact 210

limit load determined by 85

vertical acceleration 89, 172

Iliac crest 33, 137

Impact velocities 44

Inelastic analysis 65

Inertia reel 164

Inflatable restraint system 137

Instrument panel 212

Inversion tube energy absorber 112

Ischial tuberosity 35

Isolation, vibration 39

Joint deformation, seats 45, 52

Landing gear

energy absorption 40, 174

failure 43

Lap belt

anchorage 141

tiedown strap 33

Lateral load, definition 29

Lateral strength, seats 178

Limit analysis 65

Limit load, definition 36

Litter

occupant weights 196

restraint 49

strength requirements 197

testing 200

Load factor, definition 27

Load limiter, definition 36

also see energy absorber

Loads, seat design 168

Longitudinal strength, seats 168

Major impact 26

MIL-HDBK-5 51 

-

MIL-S-58095(AV) 22, 160

MIL-STD-850 33

MIL-STD-1290 22

MIL-STD-1333 33, 38

Motion envelopes 203

Net-type cushions 128

267

I.
D Continued)

Occupant weights litter design 196

seat design 166 Optical relay tube 214 Padding

application 219, 243 ductile support 244

energy absorption 230

functions 219 load-distributing 232 Plastic analysis 148

Plastic foams

materials 222

properties 219, 222

test methods 227

types 222 PROMETHEUS computer program 100 Rate sensitivity, seats 175, 179 Rebound, definition 29 Reflected shoulder straps 130 Releases, structural 54, 61 Restraint system adjusters 146, 163 aircrew 130

anchorage 67

buckle 139, 163 design loads 149 desirable features 129 dynamic overshoot 130

emergency release 139 fitting roughness 158

gunner 134 hardware materials 146 inertia reel 164

inflatable 137 lap belt anchorage 141

litter 49

reflected straps 130 requirements for seat types 48 shoulder harness anchorage 144 single-point release 129 stitching 151

tiedown strap anchorage 144 testing 180

troop 133 webbing characterietics 146, 148 weight 161

268
INDEX (Continued)

Retrofit 23

Riveted connections 51, 147

Rod pull-through tube energy absorber 121

Rolling torus energy absorber 114

Rollover 43

Rudder pedals 213

Seat

aft-facing 48

cabin 42

ceiling-mounted 45

clearance for stroke 41

component attachment 195

cushions 125

design conditions 44

dynamic response 82, 85

forward-facing 48

joint deformation 52

loads during crash 41 

I

reference point 35

side-facing, restraint for 48

strength requirements 166

stroke

computation 69

efficiency 81

minimum, crewseats 174

minimum, troop seats 178

testing requirements

dynamic 188

static 181

troop 42, 178

vibration 39

width for comfort 38

Shoulder harness anchorage 1441

Sighting systems 214

Single-point release 129

Skull fracture, prevention 210, 219, 231

SOM-LA computer program 95

Specific energy

definition 36 

-

energy absorbers 106

plastic foams 239

Static strength, definition 35

Static testing adjustment position 181

body block 183, 185

deflection measurement 186

269

,I
INDEX (Continued)

floor warping 55

litters 200

seats 181

Stitching, restraint system 151

Stopping distance 67A

Strength requirements 

*

litter 197 

k

restraint webbing 149

seats 166 

H

Stress corrosion 50 1

Strike envelopes 203

Stroke

clearance for 42

computation example 70

minimum 82, 174, 178 

i

Structural integrity, definition 35

Structural release 54, 61

Submarining

definition 32

lap belt location 141

restraint system preventing 67

Survivable accident, definition 30

Tension pulley energy absorber 123

Tension tube energy absorber 119

Testing

litter 200

restraint system 180

seat 181

Thigh tangent angle

definition 34

limits for comfort 37

Thigh tangent line, definition 34

Tiedown strap

anchorage 144

comfort 130

Tolerable deceleration 89

Tolerance,- human, definition 31

Transmissibility, definit-ion 29 

.

a

Troop seat energy absorbers 110

Tube flaring energy absorber 121

Upward load, definition 29 i

Upward strength, seats 178

Variable-load energy absorber 92, 177

Velocities, impact 44 

I

Vibration sources 39

Visionic systems 214

Warping of attachments 52, 54, 58

.1

270

S ', • , • u, "- m•• -' • ,::W•'•i•• ;:"• '"'• ' "- . ..... :• 

. .. ......... • " -
INDEX (Continued)

Webbing

characteristics 146

energy-absorbing 159

low-elongation 150

stitching 151

strength 149

thickness 150, 151

width 150, 151

wrap radius 158

Weights equipment 167

occupant, seat 166

litter 196

Welded connections 51, 147

Wire-bending energy absorber 109

2

271
DEPARTMENT OF THE' ARMY

APPLIED TECHNOLOGY LABORATORY US ARMY RESEARCH AND TECMNOLOGY LABORAIORIES AVRADCOM

F¥)RT EUSTIS VIRGINIA 23d04

READER'S SERVICE LETTER (For use in submitting comments, recommendations, correction;, and revisions for Aircraft Crash Suri,ial Design Guide)

"F ROM:

TO: Director, Applied Technology Laboratory, US Army Research and Technology Laboratories (AVRADCOM), ATTN: DAVDL-ATL-ASV, Fort Eustis, Virginia 23604

REMARKS:

I

SIGNATURE: DATE: ._

,.1

S ....... •'• " •""-#;, .. ... ,-" ......i';.•S•
DEPARTMENT OF THE ARMY I APPLIED TECHNOLOGY LABORATORY 1 UUS ARMY RESEARCH AND TECHNOLOGY LABORATORIES ,AVRADCOM

FORT EUSTIS VIRGINIA 23604

READER'S SERVICE LETTER

(For use in submitting comments, recommendations, corrections, and

revisions for Aircraft Crash Survival Design Guide)

o F ROM: I

TO: Director, Applied Technology Laboratory, US Army Research and Technology

Laboratories (AVRADCOM), ATTN: DAVDL-ATL-ASV, Fort Eustis, Virginia

23604

REMARKS:

S

I

I

4

I
A IrII!

DEPARTMENT OF THE ARMY

-APPI. IEO TEýHNOLOGV I.ABORA T 014

US AHMV HILESA4CH AND0 TECHNOLOGY LASORAIOHIES AVRADCOM

FORT EUISTIS VIRGINIA 23604

READER'S SERVICE LETTER

(For use in submitting comments, recommendations, corrections, and

revisions for Aircraft Crash Surviria Design Guide)

FROM:

TO: Director, Applied Technology Laboratory, US Army Research and Technology

Laboratories (AVRADCOM). ATTN: DAVDL-ATL-ASV, Fort Eustis, Virginia

23604

REMARKS:

SIGN'ATURE: DATE: