. . . . . . . . . . . . . . . . . . . - . 21~',.. fir'-r~~ m

USARTL-TR-79-22E

AIRCRAFT CRASH SURVIVAL DESIGN GUIDE

VOLUME V -AIRCRAFT POSTCRASH SURVIVAL

8IMULA INC.

2223 SOUTH 48TH STREET

TEMPE, ARIZONA 85282

JANUARY 1980

0 FINAL REPORT

~ PREPREDlFO

FOEPRT D FOTIVRGII 30

Approved for public release; distribution unlimited._

b~qJ
APPLIED TECHNOLOGY LABORATORY POSITION STATEMENT r,

rhis 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. Subsequent revisions

were published as USAAVLABS Technical Report 70-22 and

USAAMRDL Technical Report 71-22. This current editicn consists of a consolidation of design criteria, concepts, and

analytical techniques developed through research programs '.1

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.

Th~s document has been coordinated with USAAVRADCOM, the

U. S. Army Safety Center, the U. S. Army Aeromedical Re- .... . search 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. E. Bywaters, W. J. Nolan, and H. W. Holland

of the Safety and Survivability Technical Area, Aeronautical

Systems DivisioA, Applied Technology Laboratory.

Comments or suggestions pertaining to this Design Guide will

be welcomed by this Laboratory.

DISCLAIMERS

The findings in this er arn not to be construed a an official Department of the Army position uh55 so dosinated by 6ther -,uilhrlsed docu monts. "•'

Vft" Government drwt•Ig, specifictions. or other data we used for any purpose other than In oonn•ction

with a dafinitaly related G&overnment procuremewt operation, the United States Government thereby Incau no arep iblity nor any oblgation whaetoevme and the fact that the Government may have. formuleud, furnished, or In any vIy aupplied the aild drswlr spllpec atiaon, or other date Is not to be r r by Implication or otherwls mes In any manner licensing the holdw or any other person or corporation, or conveying any rights or permlisson, to manufacture, use, or a-1l any patented Invention that may in any vy be reletd thereto.

Trade name cited In this reort do not constitute an official endoremanet or approval of the use of such owmmerdial hardwore or softwa.

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 Design Guide (formerly USAAMRDL Technical Report 71-22). All

previous editions are obsolete and should be destroyed.

;-t'•
Unclassified &RCUFAITY CL.AULFICATION OF THIS$ PAGE ("~On Doe 9ta Etrod)_________________

ffEPORT DOCUMENTATION PAGE BEFORE COMPLETlING FORM

M/R W~kM 2OVT ACCESSION NO. 3. RECIPIENT'S CATALOG NUMBERN

JSPART TR-79--22E __

4. TITLE _(ad. dubttttej -E~ fat FPA_ &ID OI"

AIRCRAFT QMSH .WRVIVAL .;ýSIGN.WLIDE, NLE

qC LUME V.inrcWo~t Postirash A~vijval _6p*7 e

f. TR--7823 /ý

7.~~~~1 .- AUtOtd ,.-CNTRACT Olt GRANT NUMUER(l)

tO N. 13./JohnsonMS. H.1 Robertson /4 DAAJý2-77-C-#'2l

4. PERFORINGGANIZATION NAME ANO AGGROESS i0. PROGRAM ELEMENT. PROJECT, TASK

Simula Inc.ARAAWRUNTNMII

2223 South 48th Street I ~ 62209AH7 ______

00 ~ 7

Tempe, Arizona 85282 F9 0_______1w

Ii. CONTROLLING OFFIC NAME AND ADDRESS A., Repopu.T.BaX._.

Applied Technology Laboratory, U. S. Arm V&7

Research and Technology Laboratories AWMS15A-OF INAGES,-

(AVk'ADCOM), Fort Eustis, Virginia 23604 224_________

14. MONITORING AGENCY NAME At AODDRESS((I different from CcoiIrollind Office) 15. SECURITY CLASS, (of thist. ,eool)

- . -"Ia, DOECLASIIP ICATION/DOWNORADING

IS, DIST"IUTO ITTE T (of this Rleport)

Approved for public release; distribution unlimited.

I?, DISTRIBUTION STATr~MENT (faith* obatl roc re t from Report)

1S. SUPPLEMENTARY NOTES

Volume V of a five-volume report

IS. KEY WRODS (Continue an reverse side if nec.esary and Identify by block number)

Aircraft. Design Guide Crash Locator Beacons

Crashworthiness Emergency Escape

Postcrash Survival Aircraft Interior Materials

Crashworthy Fuel Systems Ditching

Postorash Fire Ignition Source Control

!0. A3SISIIACT (Crcatima en rever. t~ " f 7e.euymd Idenfifby r block onumber)

This five-volume document has been assembled to assist design

engineers in understanding the problems associated with the

development of crashworthy U. S. Army aircraft. Contained herein

are not only a collection of available information and data pertinent to aircraft crashworthiness but suggested design conditions

and criteria as well.. The five volumes of the Aircraf+-'-rash

.-Survival Design Guide cover the following topics:

DO U 47 EDITON OF 0 NOV 69 15O1SOLVTE Unclassified

SEC~uRITY CLASSIFICATIOw or THIS PAGE ('WIrer Dots Entered)
Unclassified

!KCUIPIgr Y GLA:SSIFICATION Of THIS PAGIZ('WIIU fat. Enlered)

20. (contd)

Th>-Volume I - Design criteria and checklists,

-.Volume III -:Aircraft Structural Crashworthiness,

01Volume IV -ý Aircraft Cratsh Renvriromnts and Huantoleranc

"'Volme V-' Aircraft Postcrash Survival'

This volume (Volume V) contains information on the aircraft post-I

crash environment and design techniques that can be used to reduce

postcrash~ hazards. Topics include the postcrash fire environment,

crashworthiy fuel systems, ignition source control, fire behavior

of interip~r materials, ditching survival, emergency escape, and ~ crash loc tor beacons.

ow -

SECURITY CLASSIFICATION OF THIS PAGIEtfhn Pata Eanteted)
PREFACE

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

search 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 1L162209AH76. This guide is a revision of USAAMRDL

-.Technical Report 71-22, kaishSurvival Des6inGude, ubIighe6.d.

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. U. S. Air Force, Federal Aviation Administration, NASA, and S. Navy for their research in crash survival. Appreciation

Sis extended to the following organizations for providing accident case histories leading to the establishment of the impact

conditions in aircraft accidents:

• U. S. Army Safety Center, Fort Rucker, Alabama.

. 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 editions of the Crash Survival Design

Guide are most noteworthy:4

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

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

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

"Volume V has been coauthored by N. B. Johnson and S. H. Robertson. Dr. J. W. Turnbow provided valuable assistance with his

recommendations and review of this document. Appreciation is

also extended to the staff members of Simula Inc. and the Crash

Researcn Institute for their contributions.

i l 3' .t7 1
II

TABLE OF CONTENTS

Section Title Page

PREFACE . . ........... ............. 3

LIST OF ILLUSTRATIONS ... ......... ....... 8

LIST OF TABLES. . . . . . . . . .............. 12

INTRODUCTION. . . . . . . . . . . . . . .*. . . . . . 15

CHAPTER 1. BACKGROUND DISCUSSION. .......... 18

CHAPTER 2. DEFINITIONS. . . . . . . . . . . . . . . 20

2.1 GENERAL TERMS ....... ........ . 20

2.2 FUEL, OIL, AND HYDRAULIC SYSTEM TERMS. 21

2.3 IGNITION SOURCE CONTROL TERMS.... . . 21

2.4 INTERIOR MATERIALS SELECTION TERMS . . . 22

2.5 DITCHING AND EMERGENCY ESCAPE TERMS. . . 23

CHAPTER 3. POSTCRASH FIRE ...... . . . . .. 26

3.1 INTRODUCTION . . . ............ 26 4 3.2 POSTCRASH FIRE ENVIRONMENT . . . . . . . 26

3.2.1 Heat . . . . . . . . . . . . . . . . . 27

3.2.2 Smoke and Toxic Gases. . . . . . . . .. 30

3.3 HUMAN SURVIVAL AND ESCAPE ....... ... 33

3.3.1 Human Tolerance to Heat. . ........ . 33

3.3.2 Human Tolerance to Toxic Gases . . . " 37

3.3.3 Human Tolerance to Miscellaneous Fire

Factors. . .. . ............... .41

CHAPTER 4. POSTCRASH FIRE PROTECTION. . . . . . . . 43

4.1 INTRODUCTION . . ............ 43 4

4.2 FUEL CONTAINMENT . . . . ........... 43

4.2.1 Fuel Tanks ..... .......... . . . . 44

4.2.2 Fuel Lines . ........... . . 57

4.2.3 Supportive Components. . ..... 66

4.2.4 Fuel System Full-Scale Crash Test . . 77

4.3 OIL AND HYDRAULIC FLUID CONTAINMENT. 77

4.4 FUEL MODIFICATION ..... .......... . 79

4.5 IGNITION SOURCE CONTROL. . . . . . . 81

4.5.1 Electrical Sources . . .......... 81

4..5.2 Engine . . . . . . . . . . . . .. . . . 84

4.5.3 Heaters. . . . . . . ........ .. 97

4.5.4 Sparks . . . . . . . . . . . . . . . . . 97

4.5.5 Initiating Systems . ...... . . 99

li

5

£O5DULO PAM XLNMWrM 1IIJD
TABLE OF CONTENTS (CONTD)

Section Title Pav

CHAPTER 5. INTERIOR MATERIALS.... . ......... . . 103

5.1 INTRODUCTION . . . . . . . . . . . . . . 103

5.2 FIRE BEHAVIOR OF MATERIALS . . . . . . . 104

5.2.1 Flammability . . . . . . . . . . . . . 104 5.2.2 Smoke. . . .. . ........... 110

5.2.3 Toxic Gases . . . . . . . . . . . . . 113

5.3 MATERIAL TESTING . . . . . . . . . . . . 115

5.3.1 Latoratory Testing. . . . . . . . . . 115

5.3.2 Large-Scale Testing. . . . . ... .. 119

5.4 SELECTED MATERIAL PROPERTIES . . .. . . 120

5.4 1 Seat Cushion Foams . . . . . . . . . . . 121

5.4.2 Upholstery and Other Fabrics . . . . . . 127

5.4.3 Structural Components. . . . . . . . . . 1.29

5.4.4 Aircraft Insulation. . . . . . . . . . . 129

5.5 MATERIAL SELECTION CRITERIA. . . . . ... 135

5.5.1 General Considerations . . . . . . . . . 135

5..5.2 Flammability Test Criteria . . . . . . . 136

5.5.3 Evaluation of Materials. . . . . . . . . 139

CHAPTER 6. DITCHING PROVISIONS. . . . . . . . . . . 145

6.1 INTRODUCTION . . . . . . . . . . . . . . 145

6.2 THE DITCHING ENVIRONMENT . . . . . . . . 145

6.2.1 Aircraft Configuration and Survivability Characteristics . . . a . . . . . . 145

6.2.2 Underwater Escape . . . . . .6 . . . . 147

6.3 EMERGENCY EGRESS OPENINGS. . . . . . . . 148

6.3.1 General Provisions . . . . . . . . . . . 148

6.3.2 Explosively Cut Exits. . .... ..... 149

6.4 UNDERWATER EMERGENCY LICHTING ..... 149

6.4.1 Underwater Visibility. . . ....... 149

6.4.2 Emergency Lighting Requirements. 150

6.5 AIRCRAFT FLOTATION SYSTEMS . ...... .. 156

6.5.1 Sponsons . ' . . . . . . . . 156

6.5.2 Flotation Bags . . . . . . ....... 157 A

6.6 DITCHING EQUIPMENT . . . . ....... 158

CHAPTER 7. EMERGENCY ESCAPE PROVISIONS. . . . . . . 160

7.1 INTRODUCTION ............. 160

7.2 EMERGENCY EXITS.. . . • ...... 160

7.2.1 Emergency Exit Requirements. .... 160

7.2.2 Types of Exits . .. ... .. ............ 161

7.2.3 Size of Exits. . . . . 161

7.2.4 Number of Exits. . . .. ... .... . 162

6

.~... . . .--- --.. ,,
TABLE OF CONTENTS (CONTD)

Section Title Page

7.2.5 Location of Exits ............... ... 164

7.2.6 Exit Operation ............ 165

7.2.7 Explosively Created Exits. ........ ... 170

7.2.8 Access to Exits .... ......... . . .. 174

7.3 EMERGENCY LIGHTING .... . ...... ... 175

7.3.1 Interior Emergency Lighting. ...... ... 175

7.3.2 Emergency Exit Lights. .......... 176

7.3.3 Exterior Emergency Lighting. . . . . . 178

7.3.4 Structural Considerations ............ 178

7.3.5 Emergency Lighting Power Source. . 180

7.3.6 Actuation of Emergency Lighting Units. 180

7.4 MARKING OF EMERGENCY EXITS . .... 181

7.4.1 Internal Identification of Exits . . . . 181

7.4.2 External Identification of Exits .... 182

7.4.3 Identification of Secondary Openings . . 182

7.4.4 Marking Instructions for Exit

Operations ...... . .. . . .. . . 182

7.5 CREW CHIEF STATIONS. . . . . . . . . . . 183

CHAPTER 8. CRASH LOCATOR BEACONS. . . . . . . . . . 185

8.1 INTRODUCTION........... . . . .... . 185

S8.2 TYPES OF LOCATOR BEACONS . . . . ... 185

8.2.1 Automatic and Manual Activation .. 185

8.2.2 Fixed, Portable, and Deployable

Equipment ..... .... .... . .... .. 185

8.3 COMPONENTS OF LOCATOR BEACONS. . .... 186

8.3.1 Crash Sensors. .............. 186 1 8.3.2 Transmitter ..... ......... . 197

8.3.3 Antenna.............. . . . . . . . . 198

8.3.4 Power Supply. . .... ...... . . .. . 198

8.3.5 Electrical Wiring.. ............. ... 199

REFERENCES .......... ...................... ... 200

BIBLIOGRAPHY .......... ..................... ... 210

APPENDIX A ...... ...................... ... 217

INDEX ......................... . . 219

° ' ' ' ' ° ' • ' ' 7 ' ' ' '-2' .1 1
LIST OF ILLUSTRATIONS

Figure Pag

1 Average recorded ambient and radiant

temperatures in large, crashed, burning, passenger/cargo-carrying, fixedwing aircraft. . ........ .*. . . . 28

2 Recorded ambient temperature range in

the cabin areas of large, burning,

passenger/cargo-carrying helicopters . . 29

3 Aircraft skin melting time based on

I gross weight . . . . . . . . . . . . . . 30

4 Average recorded CO concentrations and

calculated COHb levels in large,

crashed, burning, passenger/cargocarying fixed-wing aircraft. . . . ... 31

5 Average recorded CO concentrations and

calculated COHb levels in large,

crashed, burning, passenger/cargocarrying helicopters . . . . . . . . . . 32

6 The hemisphere of radiant heat

concept ...... .............. . . . . 35

7 Pain threshold time as a function of

temperature of radiant heat source . . . 35

8 Human tolerance to ambient air

temperatures ..... .............. .. 37

9i Physiological effects of various

carboxyhemoglobin percentages ...... .. 38

10 Probable CO uptake in humans exposed to

various air concentrations of COHb . . 39

11 Tank shapes ..... ............... ... 46

12 Resistance of materials to tearing... 49

13 Slit test for obtaining resistance to

tear propagation .... ............ .. 50

14 Drop test fixture .... ............ .. 52

8

;! ' A
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

15 Typical setup for dynamic testing of

fuel cell fitting pullout......... 53

16 Typical fuel cell fitting pullout

following dynamic test . . . . . . . . . 53

17 High-strength fitting retention

techniques . ... . .. . .... . . . . . 55

18 Sample frangible attachment separation

load calculation . . . . . . . . . . . . 56

19 Specifically designed bulkhead (firewall)-to-line breakaway valve .... ...... 61

20 Modified quick-disconnect line-to-line valve . . . . . . . . . . . . . . . . . . 61

21 Typical breakaway load calculation for

in-line breakaway valve. . . . . . . . . 62

22 Standard hose fitting dimensions .... 63

23 Self-sealing hose fitting dimensions . . 64

24 Hose stabilizing with frangible

structure. . ........... . . . . . 66

25 "One-shot" self-sealing valve ...... 67

26 Cable-actuated quick-disconnect valve. . 67

K. 27 Typical cell-to-cell self-sealing

breakaway interconnect valve ...... 68

28 Typical method of breakaway load

calculation for fuel tank-to-line

breakaway valve ............ . . . . . . 69

29 Vent valve . . . . ............... 71

30 Frangible attachment of air-driven

boost pump to airframe structure .... 73

31 Frangible filler neck installation . . . 73

A ( 9

* ~ i
i'9

LIST OF ILLUSTRATIONS (CONTD)

Figure Page

32 Check valve installed in conjunction

with frangible filler neck . . . . . . . 74

33 Frangible ring attachment for

installing fuel tank fillers . . . ... 74

34 Crashworthy fuel quantity sensor .... 75

35 Design concepts for crashworthy fuel

drain cocks. . . . . . . . . . . . . . . 76

36 Felt oil tank cover. . . . . . . . . . . 79

37 Cable cutter de-energizing method. . . . 83

38 Rotating beacon installation . . . . . . 85

39 w*P-4 ignition delay versus surface

temperature. . . . . . . . . . . . . . . 87

S40 Hot surface inerting system ........... 90

41 Flame-source inerting system ........ 91

42 Schematic diagram showing fuel shutoff

and drain system in J57 engine crashfire protection system ......... 92

43 Fuel baffle and fire curtain concept . . 94

44 External drip fence. . . . . . . . . . . 95

45 Internal drip fence. . . . . . . . . . . 95

46 Fuel flow diverters . . . . . . . . . . 96

47 Block diagram showing how initial

switches are incorporated into a

crash-fire-prevention system for

twin-engined airplane. . . . . . . . . . 101

48 Potential activating circuitry for

single-engine aircraft . . . . . . . . . 102

10

'~~~~~ , ~~~ ..•' ..."•d• . . •.'.•• ,i
LIST OF ILLUSTRATIONS (CONTD)

4. Figure Page

49 Influence of superimposed heating on

minimum oxygen concentration for various

polymer systems burning in the drivenrod configuration. . . . . . . . . . . . 107

50 Influence of superimposed heating on

flame spread rates of various

polymers . . . . . . . . . . . . . . . . 108

51 Effect of concentration of the reactive

fire retardant o,o-diethyl-n,n,-bis

(2-hydroxyethyl) aminomethylphosphonate

on light obscuration in rigid-urethane

foams .c.r.t.o . . .r.g.d.u.e . foa. . ... 112

52 Effect of the nonreactive fire retardant

tris, 2,3-dibromopropylphosphate on light Sobscuration in rigid-urethane foams.. 112

53 Seat back temperatures during burn tests

of coated and uncoated polyurethane

foam seats ............... 123

54 Hydrogen cyanide concentrations from

burn tests of coated and uncoated

polyurethane foams ........... 124

55 Semirigid foam insulation test

configuration . . . . . . . . 133

56 Compartment ambient temperature versus

time from fuel ignition during foam

insulation tests ...... ...... 134

57 Comparison of luminance thresholds at

two adapting levels in water and in air . . . . . . . . . . . . . . . . . . . 151

58 Values of V in Equation (3) for distances from light source to observer

with various water attenuation coefficients (a) . . . . . . . . . . . . . . 153

59 Threshold sensitivity (S) levels for

elapsed time after ditching for adaption fields (AF) from 0.35 to 350

cd/ma. . . . . . . . .v.. . . . . . 154

11,!

I' " ', , ,
LIST OF ILLUSTRATIONS (CONTD)

Figure Page

60 Tolerance time for different water

temperatures while in life rafts . . . . 159

61 Schematic of ELSIE system. ........ .. 171

62 Relative brightness of internally

illuminated aircraft exit signs at

various distances from signs . . . . . . 177

63 Light transmittance of exit signs in

clear air and in smoke at various

distances from sign. . . . . . . . . . . 179

64 Typical exit release instruction:

incorporating symbols ...... . . 184

65 Longitudinal velocity changes for

survivable rotary- and light fixedwing aircraft accidents. . . . . . . . . 189

66 Vertical velocity changes for survivable

rotary- and light fixed-wing aircraft

accidents. . . . ............... 190

67 Worst vibration loads superimposed on

severe longitudinal components of flight

loads . . . . . . . . . . . . . . . 192

68 Proposed specification for crash force

sensors in light fixed-wing aircraft . . 194

69 Proposed specification for vertical

crash force sensors in rotary-wing

aircraft . ...................... ..... 195

LIST OF TABLES

Table

1 Tolerance to selected combustion

gases. . . . . . ................. 40

2 Required minimum individual loads for

standard hose and hose-end fitting

combinations .. . . . . ........ .. 58

12

. . . . . A...............-**
LIST OF TABLES (CONTD)

Table Page

3 Required minimum individual loads for

self-sealing hose and hose-end fitting

combinations...... . . . ..*. . 59

4 Required minimum individual loads for

self-sealing hose with flanged end fittings o. . . . . . .. o. . . . .. . . 60

5 Summary of design velocities for rotaryand light fixed-wing aircraft. . . . . . 77

6 Minimum conditions under which certain

abraded metal particles will ignite. . . 98

7 Minimum autoignition temperatures (AIT)

and hot plate ignition temperatures of

sheet-type combustibles in air . . . . . 105

8 Toxic gases produced by burning aircraft

interior materials....... . . . . . . 113

9 Results of large-scale air transport

cabin fire tests using urethane or

neoprene seat cushions . . . . .. . 125

10 Typical properties of standard and

improved neoprene foam . . . . . . .. 125

S~imide, 11 Comparison of polyphosphazene, poly- and fire-retarded polyurethane

foam properties . ...... . . . .. 126

12 Flammability properties of various

fabric materials . . . . . . . . . . . . 128

13 Preliminary properties of candidate

•i. compression molding materials. . . . . . 130

14 Preliminary properties of candidate

thermoformed materials ...... . . . 131

15 Flammability appraisal of materials used

in commercial aircraft . . . . . . . . . 140

16 Survival rate versus helicopter

flotation time . . . . . . . . . ... 147

13

1,•• k.' "•,Y ,•••I "
LIST OF TABLES (COND)

Table Page

17 Representative attenuation coefficients

(a) for various water sources. . . . . . 151

18 Examples of values for minimum luminance

(L) in cd/m2 for escape hatch lights at

distances to level of 350 cd/m2 ..... 154

19 Vibration spectrum characteristics

based on Radio Technical Commission

for Aeronautics, DO-147 requirements . . 191

20 Vibration spectrum characteristics

based on MIL-STD-810B requirements . . . 193

21 Summary of design conditions for crash

locator beacons. . . . . . ........... 196

14

D '~
INTRODUCTION

For many years, emphasis in aircraft accident investigation

was placed on finding the cause of the accident. Very little

effort was expended in the crash survival aspects of aviation

safety. However, it became apparent through detailed studies

of accident investigation reports that large improvements in

crash survival could be made if consideration were given in

the initial aircraft design to the following general survivability factors:

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 becoming missiles during a crash sequence.

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 between crash forces,

structural failures, crash fires, and injuries. A series of

reports covering this effort was prepared and distributed by

the U. S. Army, beginning in 1959. 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

*Now the Applied Technology Laboratory, Research and Technology Laboratories of the U. S. Army Aviation Research and Development Command (AVRADCOQI).

15

. *..I-,,
engineers 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 Soa1.

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

research in crashworthiness technology. The last revision

of TR 71-22 was the basis for the criteria contained in the

Army's crashworthiness military standard, 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 Aircraft Crash Survival Design Guide is published in five volumes. Volume titles and general

subaects included in each volume are as follows:

Volume I - Design Criteria and Checklists

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,

accident information retrieval.

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 1972.

16

I ; ;h
Volume IV - Aircraft Seats, Restraints, and Litters

l Operational and crash environment, energy attenuation,

seat design, litter requirements, restraint system design, occupant/restraint system/seat modeling.

Volume V - Aircraft Postcrash Survival

Postcrash fire, ditching, emergency escape, crash locator beacons.

This volume (Volume V) contains information on the aircraft

postcrash environment and design techniques that can be used

to reduce postcrash hazards. It contains a great deal of background information, including data from such sources as fullscale aircraft burn tests, laboratory materials testing, and

research and development programs in aircraft fuel systems.

Chapter 1 presents a general discussion of designing for crashworthiness. Chapter 2 contains definitions of terms pertinent

to the volume. Chapter 3 describes the postcrash fire environment and relates this environment to human tolerance data

in the areas of heat, smoke, and toxic gases. Chapter 4 discusses methods of preventing postcrash fires by containing

flammable fluids in crashworthy fuel, oil, and hydraulic systems, modifying fuel properties to reduce crash-induced fuel

misting, and controlling potential ignition sources. Chapter 5 discusses the fire behavior of interior materials and

presents data on material flammability tests and selected material properties. Chapter 6 describes the ditching environment and provisions that can be incorporated into the aircraft

design to increase ditching survival. Chapter 7 presents design requirements for emergency escape exits and emergency

lighting, and Chapter 8 discusses crash locator beacons.

'II

( 17

-", 1 7

OI..IL
1. BACKGROUND DISCUSSION

This volume specifically addresses the hazards that exist in

the postcrash phase of U. S. Army aircraft accidents and presents aircraft design criteria that will, if followed, eliminate or reduce the serious consequences of these hazards.

Designing for postcrash safety is only a part of the larger

effort of designing the entire aircraft for crashworthiness.

The overall objective of designing for crashworthiness is to

eliminate unnecessary injuries and fatalities in relatively

mild impacts. 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 an

extremely wise investment. The outstanding success of the

crashworthy fuel systems in almost entirely eliminating thermal fatalities and injuries in U. S. Army helicopter accidents

provides a concrete example of the benefits that can be obtained through crashworthy design. Consequently, new generation aircraft are being procured to rather stringent crashworthy requirements.

The original edition of this design guide dealt primarily with

modifications that could be made to existing aircraft to increase their crashworthiness. Now, two approaches to improving

aircraft crashworthiness are open. The first approach is to I influence the design of new aircraft, and the second is to improve the crashworthiness of existing aircraft. Obviously,

much higher levels of crashworthiness can be achieved in the design and development of new aircraft if crashworthiness is

considered from the beginning. This is being accomplished at

the present time through the use of procurement packages that

include pertinent specifications that require certain levels

of crashworthiness for various subsystems as well as for the

entire aircraft. However, some of the available potential is

still being lost due to the historical approach used in designing aircraft. That is, the basic aircraft is designed leaving

space and providing attachment provisions for subsystems.

Later, when the subsystems are designed, their designs are limited by the previously established, somewhat arbitrary boundary

conditions. The boundary conditions may unnecessarily limit

the performance of the subsystems. The better approach is to

design all systems and subsystems at the same time, at least

preliminarily. This enables subsystem considerations to affect

the larger systems. This systems approach will produce a more

nearly optimum vehicle.

The same principles for improving crashworthiness can be applied to the retrofit of existing aircraft; however, the "castin-concrete" status of existing production structure is a more

'4' 18

FLL
costly and difficult obstacle to overcome. When crashworthiness features must be included through retrofit, the level that

can be achieved is usually reduced. Even in retrofit situations, however, the overall objective can be met; i.e., occupant protection can be maximized to eliminate unnecessary injuries.

In earlier editions of the Design Guide, the requirements to 4 provide occupant protection in crashes up to and including the

severity of the 95th-percentile survivable crash pulse was expressed. With the deployment of aircraft designed for crash

safety, the link to the 95th-percentile survivable crash pulse has been dropped, and the recommended design environment is

simply presented as the design pulse. Obviously, the severity of a 95th-percentile survivable crash pulse will be much

greater for the new aircraft than for aircraft having no crashworthy requirements placed upon them during their development.

The extent of the crash protection provided to the occupant

cannot indefinitely continue to be linked to the survivability

of the crash as improved crashworthiness increases the severity of the survivable crash, producing a never-ending increase in

the level of crashworthiness at the expense of aircraft performance. The crashworthiness levels recommended herein are felt

to be a near optimum mix of requirements including considerations of cost, weight, and performance. The crash environments

selected for design purposes in this volume are identical to

the historical 95th-percentile survivable crash pulses.

Also, in earlier editions of the Design Guide, information was

provided on design of fixed-wing transport aircraft. However,

considering the volume of new information on crashworthy design

and the need to ensure that the size of this document remains

within reasonable limits, only the primary aircraft in the Army

inventory are considered. Therefore, information given herein

is intended to apply to rotary-wing aircraft and light fixedwing aircraft, defined by a mission gross weight of 12,500 lb

or less.

41

I

S ( 19

. 1

.....................................................................
2. DEFINITIONS

2.1 GENERAL TERMS

o The Term "G"

The ratio of a particular acceleration (a negative

acceleration may be referred to as a deceleration)

to the acceleration due to gravitational attraction

at seat level (32.2 ft/sec2). With respect to the

crash environment, unless otherwise specified, all

acceleration values (G) are those at a point approximately at the center of the floor of the fuselage. In accordance with common practice, 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/sec3 . As a result, crash forces can be thought of in terms

of multiples of the weight of objects being accelerated. Therefore, in keeping with common practice,

the term G is used in this document to define accelerations or forces.

e Static Strength

The maximum static load that can be sustained by a structure, often expressed in terms of acceleration (G) of a given mass or, in other words, a load factor.

o Load Factor

A factor that when multiplied by a weight produces a

force used to establish static strength. Load.factor

is expressed in units of G.

* Forward Load

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

axis.

s Aftward Load

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

axis.

* Lateral Load

Loading in a direction parallel to the lateral

(pitch) axis of the aircraft.

20
e Downward Load

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

SUpward Load

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

* Velocity Change (AV)

The decrease in velocity of the airframe during the

ma3or impact, expressed in feet per second. The ma- jor impact is -the one in which the highest forces--re incurred, not necessarily the initial impact.

2.2 FUEL, OIL, AND HYDRAULIC SYSTEM TERMS

* Boost Pump

A fuel pump installed in the tank of an aircraft to

supply the main (usually engine-driven) fuel pump

with sufficiently high inlet pressure to meet net

positive suction head (NPSH) requirements under all

flight conditions.

* Frangible Attachment

An attachment possessinq a part that is constructed

to fail at a predetermined location and/or load.

* Fuel Valve

Any valve, other than a self-sealing breakaway valve,

contained in the fuel supply system, such as fuel

shutoff valves, check valves, etc.

0 Self--sealing Breakaway Valve

A fluid-carrying line or tank connection that will A

separate at a predetermined load and seal at both

ends so that an absolute minimum of fluid is lost.

2.3 IGNITION SOURCE CONTROL TERMS

* Fire Curtain

A baffle made of fire-resistant material that is used

to prevent spilled flammable fluids and/or flames

from reaching ignition sources or occupiable areas.

.. 2 
- Fire-Resistant Material

Material able to resist flame penetration for 5 min

when subjected to 2000OF flame and still be able to

meet its intended function.

e Firewall

A partition capable of withstanding 2000 0 F flame over

an area of 5 sq in. for a period of 15 min without

flame penetration.

* Flammable Fluid

Any fluid that ignites readily in air, such as hydrocarbon fuels and lubricants.

0 Flow Diverter

A physical barrier that interrupts or diverts the

flow of a liquid.

0 Ignition Temperature

The lowest temperature at which a flammable mixture

will ignite when introduced into a specific set of

circumstances.

e Inerting

The rendering of an aircraft system or the atmosphere

surrounding the system incapable of supporting combustion.

2.4 INTERIOR MATERIALS SELECTION TERMS

9 Autoignition Temperature

The lowest temperature at which a flammable substance

will ignite without the application of an outside

ignition source, such as flames or sparks.

* Flame Propagation Index (Is)

A number calculated by combining two factors derived

from the ASTM E 162 radiant panel test for material

flammability. One factor is derived from the rate

of progress of the flame front and the other is derived from the rate of heat liberated by the mater- ial under test.

22

..................................
*Flame ResistantI

Material that is self-extinguishing after removal of

a flame.

* Flashover

The sudden spread of flame throughout an area due to

ignition of combustible vapors that are heated to

their flash point.

* Flash Point

The lowest temperature at which vapors above a combustible substance ignite in air when exposed to

flame.

9 Intumescent Paint

A paint that swells and chars when exposed to flames.

a Optical Density (Ds)

The optical density is defined by the relationship

I . 100

Ds log T

where T is the percent of light transmission through

a medium (e.g., air, smoke, etc.).

2.5 DITCHING AND EMERGENCY ESCAPE TERMS

* Brightness

The luminous flux emitted per unit of emissive area

as projected on a plane normal to the line of sight.

Measured in foot-lamberts.

* Candela (cd)

A unit of luminous intensity equal to 1/60 of the

luminous intensity of I square centimeter of a blackbody surface at the solidification temperature of platinum. Also called candle or new candle.

23

W
I

9 Class A Exit

A door, hatch, canopy, or other exit closure intended

primarily for normal entry and exit.

o Class B Exit

A door, hatch, or other exit closure intended primarily for service or logistic purposes (e.g., cargo

hatches and rear loading ramps or clamshell doors).

* Class C Exit

A window, door, hatch, or other exit closure intended

primarily for emergency evacuation.

o Cockpit Enclosure

That portion of the airframe that encloses the pilot,

copilot, or other flight crew members. An aircraft

may have multiple cockpits, or the cockpit may be

physically integrated with the troop/passenger section.

o Ditching

The landing of an aircraft on water with the intention of abandoning it.

e Emergency Lighting

Illumination required for emergency evacuation and

rescue when normal illumination is not available.

4 Exit Closure

A window, door, hatch, canopy, or other device used

to close, fill, or occupy an exit opening.

A , Exit Opening

An opening provided in aircraft structure to facilitate either normal or emergency exit and entry.

a Exit Release Handle

The primary handle, lever, or latch used to open or

jettison the exit closure from the fuselage to permit

emergency evacuation.

24

iL4~~~~ i'.. ' " ... i ,... i'" . -; • ,-
S Foot-candle (fc)

A unit of illuminance on a surface that is everywhere

1 foot from a uniform point source of light of 1 candela.

0 Foot-lambert (fL)

A unit of photometric brightness or luminous intensity per unit emissive area of a surface in a given direction. One foot-lambert is equal to 1/r candela

per square foot.

* Illumination

The luminous flux per unit area on an intercepting

surface at any given point. Measured in foot-candles.

.3t

I

25

:1

La.'i
3. POSTCRASH FIRE

3.1 INTRODUCTION

Historically, studies of accident records have indicated that

a high percentage of fatalities occur in accidents involving

postcrash fire. During the past 10 years, however, the pattern has changed dramatically. The accident records now indicate two distinct patterns. For aircraft not equipped with

crashworthy fuel systems, the statistical records remain essentially unchanged. For aircraft containing crashworthy fuel

systems, the fire death and injury rates have been reduced to

nearly zero.

The postcrash fire environment associated with the aircraft

not containing a crashworthy fuel system consists of a combination of marny interacting hazards. The total fire threat to

the occupant depends upon the magnitude of these hazards combined with the human tolerance limits to each hazard. This

chapter describes the postcrash fire environment and discusses

human tolerance to heat, toxic gases, and other ha'ards that

greatly affect human survival in a postcrash fire.

3.2 POSTCRASH FIRE ENVIRONMENT

The postcrash fire environment has been extensively studied in

test programs as well as in actual crashes by various research

organizations including NACA (prior to becoming NASA), NASA,

FAA, AvSER, the U. S. Air Force, and the U. S. Army. During

some of the test programs, aircraft were crashed and allowed to

burn, with data being accumulated during the entire sequence.

In other test programs, previously crashed aircraft were instrumented and burned. In addition to full-scale tests, many

studies have been performed with various components and mockups, and researchers have studied actual aircraft crashes in

which occupants were exposed to the postcrash fire environment.

From these overall studies, the most significant factors influencing survivability in postcrash fires have emerged.

Briefly, it has been observed that many variables can influence

the magnitude and threat of a postcrash fire. Some of the more

pertinent ones include the relative wind, the type of terrain

onto which the flammable fluid has drained, the fuel distribution, the location of the fluid spillage within the aircraft,

the number of structural openings (designed or crash produced)

26

i•_ ••'• :• : )
that meter the inflowing air available for an internal .ire, : and the amount of fuel available to spill (Reference 2).

It was noted that using fuels of lower volatility (i.e., Jet A

rather than Jet B) makes little difference in the overall fire

threat once a postcrash fire has started (Reference 3). However, if the fuel is spilled in liquid form and kept in that

state, rather than being formed into a mist, the likelihood of

the less volatile fuel catching on fire is measurably reduced.

In other words, if the aircraft crashes and comes to a stop

with no fire, the chances of a fire then starting are generally less with fuels of lower volatility.

However, the factors that best describe the postcrash fire situation in terms of human survival are the heat, toxic gases, and smoke existing in or near the occupiable area.

3.2.1 Heat

A typical ambient and radiant temperature curve for large

cargo/passenger-carrying aircraft tested by NACA is presented in Figure 1. As can be seen on the chart, little temperature increase occurred until 80 sec after impact. One of the main

reasons for the delay in temperature rise was the protective

shield afforded by the fuselage. Skin burn-through averaged about 80 sec, although some burn-through times occurred before

40 sec and some occurred later. Calculated escape times based

on human tolerance to heat varied from 53 to 220 sec, with the

average escape time equal to 135 sec. (See Section 3.3.1 for

a discussion of the effect of.heat on escape times.)

An ambient temperature range typical for the burning passen- ger/cargo-carrying helicopters tested by AvSER and the U. S.

Army Aeromedical Research ýLaboratory is presented in Figure 2.

This chart shows that the temperature started to increase almost immediately after the crash. The early temperature rise

was due mainly to two factors. One was that extensive structural breakup occurred upon impact, causing openings that allowed air to be drawn in, providing oxygen for internal fires.

2. Johnson, N. B., et al., AN APPRAISAL OF THE POSTCRASH

FIRE ENVIRONMENT, Dynamic Science (The AvSER Facility);

USANLABS Technical Report 70-22-CE, U. S. Army Natick

Laboratories, Natick, Massachusetts, September 1969,

AD 699826.

3. CRC-Aviation Fuel Safety Report Task Force, AVIATION FUEL

SAFETY - 1975, CRC Project No. CA-52-74, Coordinating Research Council, Inc., New York, November 1975.

27 -:4
1000

8OO

Radiant

curve /

0 600 /

QA I

400

EA4

// Ambient

2_curve

,0 •verage escape

/time 135 sec

0 ii ----or7 -I- - -i

0 40 80 120 160 200

Time after impact, sec

A Figure 1. Average recorded ambient and radiant

temperatures in large, crashed, burning, passenger/cargo-carrying, fixedwing aircraft.

The second factor was that, in the normal configuration, the

fuselage and the fuel were located in close proximity to one

another. As a result, the fire and the occupiable area were

nearly superimposed from the start. Reference to Figure 2

shows that the average escape time for these helicopters was

in the range ot 7 to 16 sec.

28

'I **'', . 'L'U .4-1- ,
Aeromedicali•

160laboratory

data (open,' fuselage)• •i !

1400 "' I ,i '

1200 -

4 800

400

S-wEscape time

20range of

200 - ~-7 to 16 L.

S~sec i

00 5 10 15 20 25 30 3.5 40 45 50

Time, sec

Figure 2. Recorded ambient temperature range in the

cabin areas of large, burning, passenger/ i

cargo-carrying helicopters.

Full-scale fire tests on standard aluminum aircraft skin panels

show that, for a fuel fire of maximum severity and minimum skin

thickness, burn-through may occur in as little as 10 set.

Larger aircraft, which possess thicker skin panels, have burned , through in 30 to 40 sec. Figure 3 shows minimum skin me~lting •,

times based on aircraft gross weight. Escape times arp obVi-,

ously shorter for faster burn-through times. Thus, the ve'ry ' short escape times in light aircraft are due not only to the

proximity of the fuel to the occupant but also to the faster r

burn-through times of the thinner fuselage skins. "

2 9 '

I I I I i-I i 
S~50

S4J 30

20

10 3 2 4 6 104 2 4 6 105 2 4 6 106

Aircraft gross weight, lb

Figure 3. Aircraft skin melting time based on

gross weight.

Another factor that can influence burn-through time is insulation. When an aircraft skin is heated externally by a fire,

the metal skin attempts to radiate heat internally. When this

radiation is prevented or retarded by insulation, skin burnthrough occurs more rapidly.

3.2.2 Smoke and Toxic Gases

Aircraft crash fires generate large quantities of dense smoke

consisting of unburned carbon particles, ashes, and gaseous

combustion products. The hazards of smoke may be both physical I (blocking vision) and physiological (irritation of eyes and

respiratory tract, toxicity).

The rapid obscuration of vision by smoke has been reported by

many survivors of aircraft postcrash fires. In addition, several test programs have documented the generation of large

quantities of smoke during burn tests of transport cabin mockups used to evaluate aircraft interior materials. However, no

30

••5... ~~~~~~~~~I "..I2:]• 
quantitative data on smoke obscuration are available from fullscale crash tests.

Quantitative crash test data are available for carbon monoxide

(CO), the predominant toxic gas generated during crash fires.

The history of carbon monoxide levels typical of NACA's pas- senger-carrying aircraft experiments is presented in Figure 4.

It can be observed that the CO concentrations remained below

the 0.8 percent level for about 250 sec, at which time they

rapidly increased to 4 percent. This slow-to-develop situation

was due to the distribution of the fuel spillage and the pro- tective shield afforded the occupants by the fuselage. Also

plotted in Figure 4 is the cumulative carboxyhemoglobin (COHb) level that would be present in an individual exposed to this atmosphere. The escape-limiting 35 percent COHb would be

reached in approximately 3 minutes. (See Section 3.3.2 for a

discussion of the effect of carbon monoxide on escape times.)

53%

4%

2.0 _" -5_0_ 

35%.-

1.2Cli 0

o 0 i

U

20

0.4 10

OW__ 0

0 60 120 L80 240 300 360 420 480 540

Vl Time, sec

Figure 4. Average recorded CO concentrations and calculated

COHb levels in large, crashed, burning, passenger/ cargo-carrying fixed-wing aircraft.

N i! ( 31-
The CO levels typical of burning, large, passenger/cargocarrying helicopters are presented in Figure 5. It can be

seen that measurable CO levels started at about 20 sec and

that within 45 sec the CO levels had increased to 3 percent.

As with the temperatures, the rapid increase of CO was due to

the fuel distribution and to the structural breakup. The CO

concentration dissipated after 45 sec due to two factors.

First, the helicopter fuselages were nearly consumed by fire

in 45 sec; thus, they could no longer act as shells to hold

the gases ii; the area. Second, only small quantities (28 gal and 56 gal) _L fuel were used during the tests. The rapid dissipation of the CO would preclude the buildup of dangerous COHb

levels in an individual exposed to these conditions.

3.0 .. 30

4j CO

S2.5 - __- - - -25

Q)

P 2.0 120 ý

00 S1.5 15

0 0.

0 co

10 20 30 40 50 60 70 80 90 1000

SFigure 0. Average recorded CO concentrations and 0alult1

COHb levels in large, crashed, burning, passenger!

S~cargo-carrying helicopters.

•. Although carbon monoxide is produced in larger amounts than

•' any other toxic gas, large-scale burn tests show that many

%• other gases also are generated, including significant amounts

W I

32j )

. , , . " Y ., ', , : 

-. .*. 0
of hydrogen chloride (HCl) and hydrogen. cyanide (HCN) (References 4 and 5). Accurate time-concentrations of these in burning aircraft fuselages are not known at this time, but the

large quantities present after 3 to 5 minutes raise the

possibility that a gas other than carbon monoxide might, at

times, be the escape-limiting factor. In fact, the recognition

that HCN was a combustion product of many aircraft materials

prompted Civil Aeromedical Institute (CAMI) scientists to determine the HCN levels in blood specimens from victims of aircraft accidents involving postcrash fire (Reference 6). It

was definitely established that HCN was present at levels

greater than normal in the blood of several victims.

3.3 HUMAN SURVIVAL AND ESCAPE

One's ability to perform a self-initiated escape from a burning aircraft becomes hampered when one is unable to think and

act as a normal human being. The point at which the incapacitating effect occurs is called the escape limit. An occupant's

escape limit is governed by what the person feels (temperature),

breathes (toxic gases), and sees, or, in case of smoke, does

not see (escape routes, blocked exits, etc.). Human tolerance

limits define human body reaction to these factors.

3.3.1 Human Tolerance To Heat

The literature dealing with the subject of human tolerance to

heat exposure is rather extensive, but somewhat confusing and

misleading. (For the purpose of this discussion, human tolerance to heat is considered for short-term exposures, up to

15 min, rather than heat prostration-type injuries that require

a considerably longer exposure time.) Although heat tolerance

has been reliably investigated by many researchers, their reports are not always clear, especially in regard to protective

measures taken during exposures to extreme heat. The reports

by Johnson and Pesman are considered to be the best application

4. Marcy, J. F., A STUDY OF AIR TRANSPORT PASSENGER CABIN

FIRES AND MATERIALS, Federal Aviation Agency; Technical

Report FAA-ADS-44, National Aviation Agency, National

Aviation Facilities Experimental Center, Atlantic City,

New Jersey, December 1965, AD 654542.

5. Heine, D., and Brenneman, J., THE FIRE TEST RESULTS, The

Airline Pilot, Vol. 35, No. 10, October 1966, pp. 8-11, S~ 18-19.

6. Mohler, S. R., AIR CRASH SURVIVAL: INJURIES AND EVACUATION TOXIC HAZARDS, Aviation, Space, and Environmental

Medicine, January 1975, pp. 86-88.

33

• . I , , I I I I I I I I4, I,
of vcientific knowledge to the subject of human thermal tolerance luring the aircraft crash-fire environment (References 2 aETherefore, iwr- cE the material in this section has

been h. _ :e. upon those reports.

Thermal injuries occurring in aircraft crash fires can be divided into two general types: skin injury and respiratory injury.

3.3.1.1 Skin Injurv; When exposed to heat, two main factors

govern a person's survival ability. They are tolerance to pain

and the thermal level at which the exposed skin will experience

second-degree burning. References 8 and 9 state that the pain

threshold is exceeded when the human skin is heated to a temperature between 108 0F and 1130F, with normal human beings experiencing unbearable pain at skin temperatures of 1240F.

Moreover, when the skin surface temperature is raised above

1110 F, the rate of cellular destruction is more rapid than cellular repair; consequently, an accumulative injury occurs. Obviously, the extent of the injury is dependent on the heat

transferred during the exposure time.

Since the temperature values required to produce pain and skin

injury are similar, pain is a good indication that injury will

occur if the application of heat continues. Therefore, approximate escape limits can be based on extreme pain and, thus,

the occurrence of radiative second-degree burns.

To approximate the occupant escape limit as fixed by radiant

temperature, one additional factor must be considered; i.e.,

the radiating surface visible to the exposed area. A hemisphere is considered to be the maximum possible radiating space

angle (Figure 6). Figure 7 shows pain threshold times aa determined by temperature of the radiative source for several

angles of radiation. If, for example, the entire hemispheric

7. Pesman, G. J., APPRAISAL OF HAZARDS TO HUMAN SURVIVAL IN

AIRPLANE CRASH FIRES, NACA Technical Note 2996, Lewis

Flight Propulsion Laboratory, National Advisory Committee

for Aeronautics, Cleveland, Ohio, September 1953.

8. Buettner, K., Ph.D., EFFECTS OF EXTREME HEAT ON MAN, Journal of the American Medical Association, Vol. 144, Octobr

1950, pp. 732-740.

9. Moritz, A. R., M.D., et al., AN EXPLORATION OF THE CASUALTY PRODUCING ATTRIBUTES OF CONFLAGRATIONS: LOCAL AND

SYSTEMIC EFFECTS OF GENERAL CUTANEOUS EXPOSURE TO EXCESSIVE HEAT OF VARYING DURATION AND INTENSITY, Archives of

Pathology, Vol. 43, 1947, pp. 466-502.

34
* Hemisphere

Exposed surface

Figure 6. The hemisphere of radiant

P heat concept.

100- Fraction of visible 1800 ~ 8. 0. C hemisphere occupied by

radiating surface.

5.0..10

4.0 I

2.0.a-- 1300 0.

2. 0.. 7 ý ___ _ _ __

U 600

0500

1.S 2 0 4 b 0 0 0 44. 

350

0.......... .6.............*,,.. .,
surface were at an elevated temperature, Curve A (F = 1.00)

would apply. If only 50 percent of the hemispheric surface

were at such temperature, Curve B (F = 0.50) would apply. The

escape limit is independent of the distance between the individual and the radiant heat source. As an example of radiant

curve usage, assume that an individual is sitting in a crashed

aircraft that is engulfed in a fireball. For all practical

purposes, the imaginary hemisphere would be 100 percent heated;

thus, Curve A would apply. Figure 7 shows that a 20-sec escape

time will be reached when the interior aircraft walls reach a

radiant temperature of only 550*F.

Experimental data on human body tolerance to convective heat

(from hot ambient air) are much more limited than data on tolerance to radiant heat. Convective heat is the primary source

of caloric uptake at low temperatures, and severe physiological

disturbances may occur at temperatures below those required ~alone for second-degree is not sufficient burning to of determine the skin. tolerance Thus, times extreme to pain heated

ambient air, and the radiative burn curves in Figure 7 cannot

be used with ambient air temperatures.

Figure 8 (from Reference 10) is a composite of the experimental

work conducted to date on human tolerance to heated ambient

air. This curve shows that the available escape time at 400*F

would be about 20 sec. This temperature is comparable to the

respiratory level temperature of 390*F selected by NACA as discussed in the following section.

3.3.1.2 Respiratory Injury: Since occupants of burning aircraft may inhale hot gases that can inflict respiratory system

injuries, a tolerance criterion is needed. However, a thorough

knowledge of rapid incapacitation from respiratory system in-

* jury is lacking. In fact, the general knowledge concerning

this aspect of human tolerance is so limited that, for all

practical purposes, there are not enough data available to establish an escape limit threshold.

A temperature of 390OF was chosen by NACA as a threshold value

to permit a gross comparison of the relative hazards of respiratory and skin injury levels (Reference 7). The 390 0 F was

chosen since it is the highest known temperature to which a human respiratory system has been exposed without damage.

10. Pryer, A. J., and Yuill, C. H., MASS FIRE LIFE HAZARD,

Southwest Research Institute, San Antonio, Texas, September 1966.

36

.... . i' •~ ~~~ý , .'•U1i3, .I 
500

o 4 0 0 -_' 0I

300

'200 -- - -- - - -

100 -

1 2 4 6 10 20 40 60 100 200 400 600 1000

Tolerance time, sec

Figure 8. Human tolerance to ambient air temperatures.

(From Reference 10)

3.3.2 Human Tolerance to Toxic Gases

The ability to escape successfully from a burning aircraft also

depends on a person's tolerance to the many toxic gases present

during a crash fire. Of these, carbon monoxide (CO) is generally the most prevalent.

The physiological effects of various carboxyhemoglobin (COHb)

levels are shown in Figure 9. During a detailed study, NACA

established that when an aircraft occupant breathes enough CO

to cause a COHb level (the percent of CO saturation in the

blood) of 35 percent, the individual's judgment becomes impaired (Reference 11). Consequently, when a COHb level of 35

percent is reached, the occupant's self-initiated escape capability becomes limited.

11. Forbes, W. H., Sargent, F., and Roughton, F. J. W., THE

RATE OF CARBON MONOXIDE UPTAKE BY NORMAL MEN, American

Journal of Physiology, Vol. 143, April 1945.

- 37

., ,.. ,. , ., . , .- ' , ", . - ' • .-. -. , .' ,. • '*"''" , ' !• "" •, " tI " ' '
90Immediate death

.- 80

• • Uncnsciousness ,

70esiratory failure,

eah in long exposure

68 0

"du gdache, disorientation, Scollapse, fainting

0

-40'Pronounced headache,

FPfatigue, irritability,

a oimpaired judgment Shortness of breath

minor during headache moderate exertion,

No major effect except shortness of breath on ý1Noextreme -physicalfft activity

Figure 9. Physiological effects of various

carboxyhemoglobin percentages.

•:'The rate of CO uptake in humans exposed to various CO concen-

• trations in the air is shown in Figure 10. These curves were '• • derived by using the following equation from Reference 10:

.• ~COHb = K x CO x t(1

where COHb w percentage of carboxyhemoglobin formed

CO - percentage of CO in the air

t w exposure time, minutes

K - absorption constant.

38

... "1 I 
r 

II

20 

4j

U

.0.

' /Z'

W

r 40

0 WI

COHb, percent p

n Figure 10 Probable CO uptake in humans exposed to,

i i wk various air concentrations of COHb.u

The absorption constant, K, depends upon the ventilation ratefer

(volume of air inhaled per minute) of the exposed person.

Since the ventilation rat es u e 10type of work being done, the constant K is equal to 3 for persons at rest, 5 for

light activity, 8 for light work, and 11 for heavy work. NACA

has chosen a ventilation rate equal tt, that of persons engaged

in light work as approximately that %;Nich would be encounitered

in persons attempting to escape from d burning aircraft (Refer-

•;•ence 7). Therefore, a value of 8 was used for the absorption

!i constant in deriving the curves. Figure 10 shows that the ;

339

L L'-. ,,
escape-limiting 35 percent COHb level is reached when the individual breathes 3.0 percent CO for 90 sec.

Although the escape times a6 limited by CO inhalation are generally longer than those limited by thermal skin injury, CO

cannot be disregarded as a serious hazard. The relationship

between ambient temperature and CO concentration will be dependent upon the type of crash, position of the occupant in

the aircraft, slope of the impacted terrain, direction of the

wind, and availability of fire-fighting equipment. It is entirely possible, especially in larger aircraft, that an individual's escape time could be limited by the CO concentration

in the air rather than by thermal injuries.

Other gases that may limit escape times from burning aircraft

include hydrogen cyanide (HCN), hydrogen chloride (HCI), nitrogen dioxide (NO ), and many others. Approximate human tolerance limits to tie most commonly expected aircraft fire gases

are given in parts per million (p/m) in Table 1 (Reference 12).

1*

TABLE i. TOLERANCE TO SELECTED COMBUSTION GASES

Hazardous levels (p/m) for

times indicated

Combustion Gas Minutes 1/2 hr 1-2 hr 8 hr

Carbon dioxide 50,000 40,000 35,000 32,000

Carbon monoxide 3,000 1,600 800 100

Sulphur dioxide 400 150 50 8

Nitrogen dioxide 240 100 50 30

Hydrogen chloride 1,000 1,000 40 7

Hydrogen cyanide 200 100 50 2

Although there is considerable variation among researchers as

to what level of a particular gas does constitute a life hazard,

the limits given in Table 1 are typical of the ranges found.

12. FIRE SAFETY ASPECTS OF POLYMERIC MATERIALS, VOLUME 6 -

AIRCRAFT: CIVIL AND MILITARY, Publication NAMB 3186,

National Materials Ad4isory Board. National Academy of

Sciences, Washington', D. C., 1977, p. 184.

40-40
Perhaps more importantly, these data illustrate the relative

lethality of the various gases. Animal experiments have confirmed that the toxicity rankings of three of the most common gases are, in decreasing order, HCN, NO2 , and HCl (Reference 13).

The escape times based on CO inhalation must be considered to

be the maximum escape times since other toxic gases present in

the crash-fire environment are much more toxic than CO. In addition, the synergistic effects of combined gases and heat on

toxicity are not well defined, although it has been established

that heated gases or combinations of gases can be more lethal

to the human than a single cool gas. Until these synergistic effects are studied in more detail, the lethal effects of each

gas in a combination must be considered to be additive to the

lethal effects of the other gases.

3.3.3 Human Tolerance to Miscellaneous Fire Factors

Discussions with survivors of actual aircraft accidents have

indicated that there are many other factors associated with

the crash-fire situation that can affect one's ability to escape. Included are visual obstructions, eye and throat irritants, fire-blocked exits, panic, and the heat factor associated with blowing hot air.

Once openings appear in the fuselage shell surrounding the occupants during a crash fire, rapid air flow through the occupiable area can begin. (It was noted during some of the fullscale aircraft burn tests conducted by AvSER that air flow '1

through the fuselage reached speeds as high as 35 mph.) This

air flow is usually hot, turbulent, and laden with toxic gases

and debris. It can create a high startle factor in the occupants because it affects their breathing and causes them to

lose sight of the surrounding area. Particulate matter in the

smoke either blocks their vision or gets into their eyes, causing the individual to close them. Further, the smoke enters

the respiratory trsct, causing severe coughing and choking.

Panic often results.

In view of the above hazards, the question of whether it is

safer to stand up or crawl out of the aircraft is often asked.

13. Higgins, E. A., Ph.D., et a]., THE ACUTE TOXICITY OF BRIEF

EXPOSURES TO HF, HCl, NO , AND HCN SINGLY AND IN COMBINATION WITH CO, FAA Civil ieromedical Institute; Report No.

FAA-AM-71-41, Department of Transportation, Federal Aviation Administration, Office of Aviation Medicine, Washing- !• tion, D. C., November 1971, AD 735160.

I41

,v".
In the opinion of the authors, there is no safe way. The turbulence in the air flow due to seats, occupants, and other

vortex generators is so great that no safe zone exists. Lowflammability clothing, a sound knowledge of evacuation procedures, and the ability and knowledge to hold one's breath while

exiting the aircraft are the occupant's primary assets for survival. Once a fire has started, the only aircraft-related

evacuation advantages an occupant can have are properly designed and located exits, escape aisles, and emergency lighting.

42 I:,

:'",4'
4. POSTCRASH FIRE PROTECTION

4.1 INTRODUCTIO14

Postcrash fire research and accident experience with the crashworthy fuel systems have shown that: (1) improvements in

ground fire-fighting systems will provide little improvement

for chances of survival in accidents where a postcrash fire is

present; (2) a reduction of fuel spillage and ignition sources

following a crash will reduce the probability of postcrash

fire; and (3) greater emphasis on "built-in" postcrash fire

protection during the aircraft design stage will improve overall postcrash fire resistance.

This chapter presents basic design guidelines for Army aircraft

systems that will inherently resist flammable fluid spillage

and ignition during survivable accidents. The fuel containment

approach is discussed first, followed by a brief summary of

fuel modification. Ignition source control, also presented,

is applicable with all forms of spillage. However, it has not,

by itself, proved to be a practical solution to the postcrash

fire problem.

When designing aircraft fuel, hydraulic, electrical, structural, and other systems, two basic requirements must be met:

(I) each system must be highly functional from the standpoint

of operational and maintenance needs; and (2) the combined system must exhibit resistance to crash fire. These requirements

can be achieved only through a design based on careful integration of the various systems, with full consideration being

given to operational and crashworthiness requirements.

The mating of the systems that offer a fire reduction potential

as well as the required operational capabilities does not necessarily imply an overall weight or cost increase. Simplicity

in the fuel system, which is desirable from the standpoint of

requiring minimum attention from the crew, may well lead to a

more crashworthy system. By following the design suggestions

contained herein, and by thoroughly understanding the fire

problem as discussed in Chapter 3, crashworthy systems that

will be practical from the standpoint of both weight and cost

can be designed.

4.2 FUEL CONTAINMENT

The design philosophy for crash-resistant fuel systems in aircraft is based upon the need to control postcrash fire in

otherwise survivable accidents. In examining the basic elements contributing to postcrash fire, three factors emerge:

an oxidizer, a combustible agent, and an ignition source.

43
Since it is not feasible to completely control the supply of

oxygen immediately surrounding the aircraft, control is best

exerciced over the remaining two elements: the fuel and the

ignition source.

The ideal fuel system is one that completely contains its flammable fluid both during and after the accident. To accomplish

this, all components of the system must resist rupture regardless of the degree of failure of the surrounding structure.

Success of such a system depends on proper selection of materials and design techniques in each of the following areas:

0 Fuel tanks.

* Fuel lines.

e Supportive components and subsystems.

There is no single, universally adaptable fuel system for aircraft. Each aircraft man-'facturer must design his own crash- worthy system based on t he criteria presented in the follow- ing sections. (A rating method to help the designer select a

crashworthy fuel system design for his particular aircraft is

presented in Reference 14.) Although the criteria given below

are specifically applicable to new aircraft design, it also is

possible to modify existing aircraft to include most of the

crashworthy fuel system principles and components.

4.2.1 Fuel Tanks

4.2.1.1 Tank Location: The location of the flammable fluidcarrying tank in an aircraft is of considerable importance in

minimizing the postcrash fire hazard from a tank installation.

The location must be considered with respect to occupants, ig- nition sources, and probable impact areas.

Greater distance between occupants and fuel supply tends to

increase escape time in the event of a fire because it reduces the likelihood of fuel entering the occupied area. Also, the tank should be kept away from probable ignition sources. While

this is not always feasible, tanks should not be installed in

or over the engine compartment, the battery, or other primary

14. Robertson, S. H., and Turnbow, J. W., A METHOD FOR SELECTING A CRASHWORTHY FUEL SYSTEM DESIGN, paper presented at

NATO/AGARD Operational Helicopter Aviation Medicine Symposium at Fort Rucker, Alabama in May 1978, North Atlantic

Treaty Organization, Advisory Group for Aerospace and De- velopment, Npuilly-sur-Seine, France.

44

MW
ignition sources. Another important consideration is the loca- ( tion of tanks with respect to probable impact damage. Accident

histories show repeated tank ruptures and consequent fires as

a result of landing gear failures, indicating the tank's high

degree of vulnerability to damage from surrounding structures.

Locating fuel tanks under a helicopter floor poses a serious

threat because of the propensity toward accidents in near-level

flight attitude at high sinking speeds. It is obvious that

fuel tanks mounted low on the fuselage will contact the ground

early in the crash sequence and will be exposed to possible

penetrations from rocks, stumps, and other ground irregularities. Thus, a good design technique is to locate fuel tanks

higher in the structure. As much aircraft structure as possible should be allowed to crush before the tanks themselves are

exposed to direct contact with obstructions.

Fuel tanks in the wings should be located as far outboard as

possible, but not at the tips. Accident investigations have

shown that placing the tanks outboard of the engine nacelles

in multiengine aircraft is preferred to locating them inboard

of the engines. Placing the tanks in the wing tips should be

avoided because these areas are anticipated impact points.

Rleduction of fuel tank volume must be considered also. If the

fuel tank is nearly full and located in an area where considerable structural collapse occurs, the tank may be subjected to

pressures that exceed its design limit. It also may be exposed

to puncture by torn and jagged metal. Therefore, if it can be

predicted that the structure surrounding the tank may collapse

due to compressive loads during a crash, expansion areas into

which the tank and its contents may displace should be provided.

Another factor that can govern whether or not a fuel tank will

survive a given impact is the method of failure experienced by

the aircraft structure surrounding the tank. Care should be

taken to ensure that when structural failure occurs in the area

of the tank, sharp cutting surfaces, penetrating spars and longeroris, and other injurious structures are avoided or controlled.

4.2.1.2 Tank Shape: The ability of the tank to displace easily and without snagging is largely dependent on its shape. .i; ~Cylindrical or rectangular shapes appear to be best, whereas 

i

tanks with protuberances or tanks composed of several interconnecting cells (see Figure 11) are most vulnerable to rupture.

Where tanks deviate greatly from the regular cylindrical or

parallelepiped shapes, consideration should be given to the

use of separate tanks with interconnecting, self-sealing fittings. To minimize snagging and excessive concentration of

45
Preferred smooth shapes

and flatwise orientation

Avoid irregular shapes

and inside angles

Figure 11. Tank shapes.

46(

I 't Ar>
stresses, inside angles should be avoided if at all possible,

especially in the lower portions of the cell. All outside an- - gles should have a radius of at least 1 in. If possible, the

tank should be oriented so that the side with the greatest

surface area is facing the direction of probable impact.

4.2.1.3 Tank Materials: The concept of fluid containment requires materials and fabrication techniques that will maximize

the energy-absorbing ability of the fuel system. Tanks constructed in accordance with earlier military specifications for crash resistance lacked such qualities and therefore failed

under minimal severity of crash conditions. Crash-resistant

fuel system research has shown, however, that fuel tanks Tonstructed of materials possessing a high degree of cut and tear

resistance, as well as a moderate degree of elongation, can

accommodate very high impact levele without loss of fuel.

These research programs resulJted in Revision B to MIL-T-27422

for crash-resistant tanks (Reference 15).

Tanks made to the specifications of MIL-T-27422B have demonstrated an ability to hold their contents safely during the

upper-limit survivable crash. However, these demonstrations

have been conducted with fuel tanks containing less than 500

gal installed in small-to-moderate-sized airplanes and helicopters. Additional research in all aspects of fuel tank

crashworthiness should be conducted before tanks with capacities exceeding 500 to 600 gal are used.

In order to provide the reader with a better understanding of

the properties of crashworthy fuel tank materials, the following general discussion is presented.

Elongation can be obtained by tank deformation or material

stretch. The amount of fuel tank elongation actually required

is unknown. It is known, however, that fuel tanks lacking the

ability to elongate are either fairly strong (heavy) or brittle,

Both types are easily ruptured in moderate crashes. On the

other hand, crash-resistant fuel tank studies have shown that

light tanks that can readily rearrange their shape (deform/

elongate), at the same time exhibiting a high degree of cut

and tear resistance, can hold their contents during upper-limit

survivable crashes.

The amount of tensile strength a tank material should possess

also is debatable. Early attempts to define a tank material

15. Military Specification, MIL-T-27422B, TANK, FUEL, CRASH-

"RESISTANT, AIRCRAFT, Department of Defense, Washington,

D. C., 24 February 1970.

47

V, I
property in terms of tensile strength-proved unsuccessful. In

fact, crash-resistant fuel system studies showed that tanks

with. lower tensile strengths were more difficult to rupture

than'ones with higher tensile values, providing, of course,

that the tanks still exhibited a high degree of cut and tear

resistance (Reference 16).

At the time of this writing, the only reason known for a minimum tensile strength requirement is to provide enough loadcarrying capability between the tank wall and the tank fitting

to cause the fitting to pull free of the airframe structure

rather than out of the tank. This usually requires the failing of some sort of frangible fastener between the tank fitting and the airframe.

What, then, defines whether or not a tank is crashworthy? The

overall results of extensive U. S. Army-funded crashworthy

fuel system studies indicated that cut, tear, and impact resistance were the key issues. ilowever, tank shape, flexural modulus of the material, reinforcement orientation, and loading

rate sensitivity were all involved. The B revision of MIL-T27422 was prepared as a result of the U. S. Army tests, and is

the best source to date to define fuel tank crashworthiness.

The cut- and tear-resistance tests, defined in MIL-T-27422B,

are self-explanatory. The values specified have proven to be

effective in actual crashes.

The importance of a material's tear resistance is illustrated

in Figure 12. These load-deflection curves were obtained

from tear tests of 3 x 7-in. specimens containing an initial

3.-in.-long slit (see Figure 13). Figure 12 shows the load required to propagate the initial slit as a function of the displacement of the pull jaws of the test device. The area under

the curve is a measure of the energy required to completely

fail the specimen. The energy required to fail the MIL-T27422B material is almost six times that required for the

0.063-in. aluminum, although the nylon/rubber composite is

lightcr in weight than the aluminum. The composite material,

though somewhat heavier due to the nature of its construction,

far surpasses MIL-T-27422A material, both in the load necessary

to propagate the tear and in the energy required to completely

16. Robertson, S. H., and Turnbow, J. W., AIRCRAFT FUEL TANK

DESIGN CRITERIA, Aviation Safety Engineering and Research

of Plight Safety Foundation; USAAVLABS Technical Report

66-24, U. S. Army Aviation Materiel Laboratories, Fort

Eustis, Virginia, March 1966, AD 631610.

48
iI

1400 o I

0.063-in. 6061-T4 aluminum

1.200

MIL-T-2744221ý

1• 000 IA ýA

400

200 -MIL-T-27422A material,-

0 1 2 3 4 ,5

Deflection, in.

Figure 12. Resistance of materials to tearing.

4 4'

" ( 49

S 4-1. ... . . .i . . " . i
.125 2.00

radius .0

7 2.00

.187 radius 2.50

TMetal test clips

.062 _ ___

7.00

Pull

2 clips

Pull test area-_/

3.00 slit

NOTE: Dimensions in inches unless otherwise

specified. Tolerances: ±.032

Figure 13. Slit test for obtaining resistance

to tear propagation.

fail the material. Further data on these materials are available in References 16 and 17.

In order to assure that proposed tank designs have seam contitniity, proper fitting installation and placement, and other

overall crash impact resistance, a drop test requirement was

17. Cook, R. L., et al., IMPROVED CRASH-RESISTANT FUEL CELL

t[MATERIAL, Goodyear Aerospace; USAAVLABS Technical Report

67-6, U. S. Army Aviation Materiel Laboratories, Fort

"Eustis, Virginia, April 1967, AD 813165L.

50

. .

* .. ! . . .. . . . . . . . . . .. - . . . , •'"
included in the MIL-T-27422B revision. Preproduction tanks in

• both the standard 30-in. cubes and the look-alike configurations, with all openings suitably closed, are filled with water

to normal capacity (air removed) and mounted on a platform of

the design shown in Figure 14. Lightweight cord is used to

maintain the cell in normal flight attitude. The platform is

raised to a height of 65 ft, released, and allowed to drop

freely onto a nondeforming surface with the platform horizontal

(±10 degrees) at impact for rotary-wing aircraft and at an angle of 20 ± 10 degrees with the horizontal for fixed-wing aircraft. No liquid leakage is allowable following the test.

The 65-ft drop height results in a severe impact test of the

fuel tank. This provides an adequate safety margin should an

aircraft crash into rough terrain (e.g., rocks and stumps),

thereby placing localized loads on the tank. Furthermore, aircraft structures surrounding the fuel tank sometimes fail in a

manner that creates additional hazards to the tank. This factor also is accommodated in the safety margin provided by the

65-ft drop test.

Review of recent crashes reveals that fuel tanks that have been

designed to the existing criteria, including the 65-ft drop

test, aie failing and releasing their contents, with fires resulting,, in accidents at and slightly above the human survival range. This verifies the validity of the design criteria. No reduction in drop height, or of cut- and tear-resistance

values, should be allowed without first conducting a major,

long-term test program to measure and define any other requirements necessary to maintain the integrity of crashworthy fuel

tanks.

4.2.1.4 Tank Fittings: A fuel cell failure often is caused

by physical displacement of the aircraft structure in relation

to the tank. This places stress concentrations at tank attachment points such as filler necks/caps, tank outlets, boost

pumps, and drains. The tank fitting can be pulled from the

tank, tearing the tank wall. Often, if the energy levels are

sufficiently high, this tear will circumscribe the entire tank.

Until MIL-T-27422B became effective, fuel cell fittings could

be torn from standard .30 caliber self-sealing fuel tanks at

loads corresponding to about one-third the strength of the tank

wall. The new specification requires high-strength insertretention methods in keeping with the high strength of the new

fuel cell materials.

MIL-T-27422B specifies that all fuel tank fittings shall have

a pullout strength of at least 80 percent of the fuel cell wall

strength. The strength of the cell material is determined by

measuring the force required to drive the end of a 4-in. diameter rod through a 13-5/8-in. diaphragm specimen of the cell

51

."! .................... ~., ~. .
':4'

NOTE: Dimensions A and B shall not exceed cell dimensions

(when the loaded cell is in place for test) by more

than 12 in, in either direction.

¶ I

Figure 14. Drop test fixture.

I material that is supported around the perimeter. The rod has

a 1/8-in. radius which forces the end into the sides. The rod

is driven at a rate of 20 in./min.

A typical method for measuring the fitting pullout strength is

shown in Figures 15 and 16. A test sample containing a 4-in.

52

"I

1'''
Figure 15. Typical setup for dynamic testing

of fuel cell fitting pullout.

anl/itig

test speci~men

Figure 16. Typical fuel cell fitting pullout

following dynamic test.

A 53

.......................................
outside diameter fitting is fabricatid of the tank material

using the same fitting material and attaching methods used on

full-size production tanks. A 200-lb weight is attached to the

fitting as shown in Figure 15. A force transducer is located

between the fitting and the weight, as close to the fitting as

possible. The test sample is attached to a rigid drop cage,

dropped from a height of 20 ft, and decelerated in a distance

of 9 in. or less. There must be sufficient distance between

the bottom of the weight and the cage to prevent bottoming

prior to fitting pullout. The peak reading from the force

transducer is the fitting pullout strength, which must be in

excess of 80 percent of the failure load of the tank material

but need not exceed 30,000 lb.

It is desirable, as a goal, for the fuel tank fitting to have

a pullout strength equal to that of the tank wall. However,

tank manufacturers have experienced great difficulty in meeting the 80 percent retention requirement currently specified

in MIL-T-27422B. Consequently, the 80 percent value is an obvious .ompromise.

Two high-strength fitting designs that have met the 80-percent retention requirement are shown in Figure 17. Proprietary designs exhibiting strengths close to or even slightly in excess

of the fuel cell wall strength have been demonstrated recently.

Thus, it may be feasible to delete this compromise in the future.

4.2.1.5 Tank Attachments: To be crashworthy, the fuel tank

must be secured to the airframe and connecting plumbing in a

way that allows the tank to pull free of the attachments without rupturing when structural displacement occurs in a crash.

Frangible brackets or bolts can be incorporated in the attachment technique to ensure their separation at specified loads.

Frangible attachments may be designed to fail either the material itself (e.g., thin-walled hollow bolts that will fail

during crash impact) or some facet of the design (e.g., protruding flanges that bend on exposure to crash forces). Several concepts, along with their applications, are illustrated

in Section 4.2.3. The frangible attachment must be strong

enough to meet all operational and service loads of the aircraft within a reasonable margin,* but should fail at 25 to 50

*A factor of 10 isia desirable goal to ensure that inadvertent

actuation under normal operation is impossible. It is realized that this goal may not always be compatible with the 50-

percent-attachment failure load criterion; however, the service load margin should be as high as possible.

54

j , • -,
.. . . .. ..........

z Metal fittin gFierg-lock retention technique

ti

Fi. ig renh fitetiong rtehntiuen t iq

jous~ syte orpietintefe ~ fFipobb er faiurdelas

.h it

'A

, .,

iFigure 17. High-strength fitting retention techniques. '

S~percent of the minimum load required to fail the attached sys- :

tem or component. This requires carefuil analysis of the var- ious components in the fuel system for probable failure loads, ,: load pathis, and degrees of deformation. A sample breakaway • load calculation is shown in Figure 18. .: It

55
Aircraft . Tank wall

structure

Frangible bolt _ Metal tank fitting

Aircraft

structureFagbe

Flange- Shear plane

ITEM LOWEST FAILURE LOAD (LB)* FAILURE MODE

Aircraft

structure 4000 Shear

Tank fitting 3000 Pull out of

tank

Flange 5000 Shear

Frangible bolt Not more than Break

3000= 1500 (tension-shear)

Not less than

3000 750

4

*Loads may or may not be representative; values are for

explanatory purposes only.

Figure 18. Sample frangible attachment separation

load calculation.

The frangible attachments should be designed to separate efficiently in the direction of force most likely to occur during

crash impact. Crash loads, whether tension, shear, compression,

or combinations thereof, must be determined for each attachment

by analyzing the surrounding aircraft structure and probable

impact forces and directions.

56
4.2.2 Fuel Lines

4.2.2.1 Line Construction: Damaged fuel lines frequently

cause spillage in aircraft accidents. Lines often are cut by

surrounding structure or worn through by rubbing against rough

surfaces. The use of flexible hose armored with a steelbraided harness is strongly suggested in areas of anticipated

dragging or structural impingement. In systems where breakaway

valves are not provided, hoses 20 to 30 percent longer than

the minimum required hose length are desirable. This will allow the hose to shift and displace with collapsing structure,

rather than be forced to carry tensile loads. For this reason,

it is equally important that couplings and fittings be used

sparingly because of their propensity to snag and restrict the

natural ability of the hose to shift.

All fittings used in the fuel system should meet the strength

requirements of Tables 2, 3, or 4 when tested in the modes

shown. The loads are always applied through the hose with

freedom allowed for the hose to form the bend radius. Thus,

the effective moment arm for the bending tests changes primarily with the line size and secondarily as the applied load

produces changes in the bend radius. This test procedure is

much easier to mechanize than one requiring a constant moment

arm.

All fuel lines should be secured with breakaway (frangible)

attachment clips in areas where structural deformation is anticipated. When fuel lines pass through areas where extensive

displacement or complete separation is anticipated, selfsealing breakaway valves should be used. The valves may be

specifically designed for this purpose (Figure 19), or quickdisconnect valves may be modified for use (Figure 20). (See

Section 4.2.3.1 for a more complete discuL.ion of self-sealing

breakaway valves.) These valves must meet all operational and

service loads of the aircraft within a reasonable margin, but

4 they should separate at between 25 and 50 percent of the min.-

imum failure load for the weakest component in the fluidcarrying system. A sample breakaway load calculation is shown

in Figure 21.

In designing a system using line-to-line breakaway valves, one

should consider potential hazards to cross-axis shear loading

on the valve halves. While omnidirectional separation is not

an ubsolute requisite for most line-to-line valves, it is

highly desirable, and every attempt should be made to procure

omnidirectional valves if there is any possibility of crossaxis shear loading.

Figures 22 and 23 will assist the designer in determining thc

lever arms and bending irioments imposed on frangible valves or

57

..........................
TABLE 2. REQUIRED MINIMUM INDIVIDUAL LOADS FOR STANDARD

HOSE AND HOSE-END FITTING COMBINATIONS

Minimum Minimum Hose end Fitting tensile load bending load

fitting type size* Wb) (1b)

STRAIGHT -4 575 450

Tension -6 600 450

1-8 900 700

-10 1250 950

-12 1900 1050

Bending = -16 1950 1450

-20 2300 1600

-24 2350 2750

-32 3500 4000

900 ELBOW -4** 575 800

Tension - -6** 600 850

8"* 900 1250

-10 1250 575

-12 1900 675

Bending * -16 1950 1200

S-20 2300 1250

-24 2350 2025

-32 3500 3500

450 ELBOW -4** 575

Tension -6** 600 425

-8"* 900 425

-10 1250 425

-12 1900 600

Bending -16 1950 1000

-20 2300 1600

-24 2350 2400

-32 3500 3700

*Fitting size given in 1/16 in. units, i.e., -4 4/16 or

1/4 in.

**Elbow material is steel.

58

T(

,p __ _ __ _ __ _ _

- d - 1
TABLE 3. REQUIRED MINIMUM INDIVIDUAL LOADS FOR SELFSEALING HOSE AND HOSE-END FITTING COMBINATIONS

Minimum Minimum

Hose end Fitting tensile load bending load

fitting type size* (ib) (ib)

STRAIGHT -10 2000

Tension -12 3120 1050

-16 2850 1650

-20 2650 1700 Bending= -24 3850 2500

-32 2700 -

900 ELBOW -10 1950 700

Tension =-12 3400 3700

:-16 3100 4300

-20 2500 2500

Bending

"-24 3800 2500

"450 ELBOW -10 1200 450

Tension = -12 3000 800

-16 3200 1800

-20 2900 1700

Bending

-24 3850 2500

*Fitting size given in 1/16 in. units, i.e., -10 =10/16

or 5/8 in.

59

""- , , 1 '- 
TABLE 4. REQUIRED MINIMUM INDIVIDUAL LOADS FOR SELFSEALING HOSE WITH FLANGED END FITTINGS

Minimum Minimum

Hose end Fitting tensile load bending load

fitting type size* (Ib) (ib)_

STRAIGHT -12 2700 3600

Tension -16 2500 1650

-24 2800 2500

Bending =

900 ELBOW -12 2400 2950

Tension = -16 2700 1050

-24 3900 2500

Bending =

450 ELBOW -12 3100 1000

Tension = -16 2100 1350

-24 3450 2500

! 'Bending

*Fitting size given in 1/16 in. units, i.e., -12 = 12/16

or 3/4 in.

'60 60

. *.. .- .... 7., ,..',*,,,, , , . .. 
'AA

Figure 19. Specifically designed bulkhead (firewall)-

to-line breakaway valve. (Tension on

attaching hose causes valve separation.)

SAircraft structure

Retainer clamp Flexible diaphragm

Load

, Release

Quick-disconnect

valve r

Figure 20. Modified quick-disconnect line-to-line

valve. (Pull of designated hose will

cause valve separation)

61

,' .._... ,, '
Component Bulkhead

Bracket Component

Hose end -Breakaway valve

T u b e t e l n b owft r sF an g i b l e •" F l e x h o e Ss a d r

fftigitanda

ITEM LOWEST FAILURE LOAD (LB)* FAILURE MODE

Flex hose 

3000 

Tensile breakage

Flex hose 

1500 

Pull Qut of end

fitn A ecin Standard 

ANfitting

fitting 1700 Break (bending)

Standard 

AN

Tube elbow fitting 1200 Break (bending) I: ~ 

Component struc- P o ~~tural attach-Pulotf

ments 4500 structure

Breakaway valve Not more than Break at fran1200 600 gible section

2

Not less than

1200

4 3

*Loads may or may not be representative; values are for

explanatory purposes only.

Figure 21. Typical breakaway load calculation

for in-line breakaway valve.

'g,,

S~~62 

!

A '
A B

Hose Maximum Maximum

A Size (in.) -.

-4 1. 33 1.16

-5 1.38 1.18

-6 1.51 1.29

-8 1.79 1.48

-10 1.94 1.60

-12 2.01 1.70

-16 2.36 1.94

-20 2.64 2.13

-24 2.79 2.18

-32 3.16 2.45

B

Hose Maximum Maximum

ASize ý(in.) ý(in.)

-B -- -4 1.72 1.16

-5 1.83 1.18

20-6 2.00 1.29 - -8 2.17 1.48

450 -10 2.42 1.60

-12 2.79 1.70

-16 3.06 1.94

-20 3.45 2.13

-24 3.65 2.18

-32 4.26 2.45

A B

A Hose Maximum Maximum

B Size (in. (in. -4 1. 59 1, 16

+2 .-5 1.68 I.18

2Q0 -6 1.85 1.29

900 -8 2.01 1. 48

-10 2.25 1.•60

-12 2.66 1.70

-16 2.97 1.94

-20 3.38 2.13

q -24 3.59 2.18

-32 4.22 2.45

Figure 22. Standard hose fitting dimensions.

I ,IA 6 3

! i K
A A B

Hose Maximum Maximum

Size (in.) (in.)

B -10 3.66 3.16

-12 3.54 3.06

-16 3.62 3.C6

-2u 3.77 3.16

I ~24 -. 3.7 b 3.06

A A B

Hose Maximum Maximum

Size (in.) 'in,)

+20 -10 3,99 3.16 -12 4.07 3.06 45 0-16 4.19 3.06

-20 4.50 3.16

-24 4.53 3.06

A- A 13

Hose Maximum Maximum

B Size (in.) (in.)

-10 3.52 3.16

-12 3.94 3.06 ,Q0 -16 4.20 3.06

-20 4. 38 3.16

-24 4.47 3.06

Figure 23. Self-sealing hose fitting dimensions.

64

S........... ~ ~ .. ~ ~ ~ ~ - .w,•.,'',•,:,', •.. ... ,
other attdching hardware of a crashworthy fuel system. The

dimensions given are for standard hose fittings. When using

nonstandard fittings, consult the appropriate drawings.

When applying these dimensions, one hose diameter (nominal)

should be added to dimension A or B, whichever is used. One

hose diameter is added because it approximately equals the offset of the load line adjacent to the hose socket as the hose

collapses when pulled in the bending mode. Dimension A plus

the nominal hose size is to be used in lever arm determinations

when standard fittings as shown are used. Dimension B plus

the nominal hose size is to be used in determining the lever

arm when other than standard elbows are used.

For example, the lever arm of a -16 size standard 90-degree

hose fitting for self-sealing hose, from Figure 23, is 4.10 in.

plus one hose diameter, or I in. Therefore, the lever arm

length equals 4.10 + 1 - 5.10 in.

The lever arm of a -10 size standard straight hose fitting for

self-sealing hose, from the same figure, is 3.66 in. plus one

hose diameter, or .63 in. Therefore, the lever arm length

equals 3.66 + .63 = 4.29 in.

For a nonstandard fitting using -10 size self-sealing hose,

the lever arm would be 3.13 in. plus one hose diameter, .63 in.,

plus the length contributed by the nonstandard component.

Therefore, the lever arm length equals 3.16 + .63 + component

length (in inches).

4.2.2.2 Line Routing: Routing of hoses should be carefully

considered during the design stage. Fuel lines should be

routed along the heavier structural members, since those mem- bars are less likely to deform or separate in an accident. Also, it is important that hoses have a space into which they

can deform when necessary. For example, when hoses pass

through large flat-plate areas, such as bulkheads or firewalls,

the hole allowing line passage should be considerably larger

than the outside diameter of the line. Hose stabilization as

well as liquid-tight, fire-tight seals still can be maintained

if a frangible structure, such as shown in Figure 24, is used.

If design requirements limit the use of the protective measures

discussed above, full use should be made of self-sealing breakaway couplings located in areas of anticipated failures. Crossover connections, drains, and outlet lines present a special

problem since they are usually located in the lower regions of [ the tank, where they are vulnerable to impact damage. Space

and flexibility should be provided at the connections to allow

room for the lines to shift with collapsing structure. Utmost

(665

4

S• • • a
Continuous hose

Grommet J ' Rigid bulkhead

i Doubler

ring

Bulkhead

urangible

material '1 -Frangible -

mp.terial

Figure 24. Hose stabilizing with frangible structure.

consideration should be given to using self-sealing breakaway

fittings at each line-to-tank attachment point.

4.2.3 Supportive Components

supportive components play a vital role in crashworthy fuel

systems. Aside from providing a solution to specific problems,

e.g., a strainer to help clean fuel, they also must be capable

of preventing spillage in accidents with resulting forces equal

to or better than the tank strength. They must not be the weak

66
link in the system. Care must be taken during the design and

testing phase to ensure that the supportive items, some of

which are discussed below, will not fail during the crash sequence and allow spillage.

4.2.3.1 Self-Sealing Breakaway Valves: Self-sealing breakaway valves are valves designed to separate into two or more sections and seal the open ends of designated fluid-carrying passages. The openings may be in fuel/oil lines, tanks, pumps, fittings, etc. The valves fall into two general categories: the "one-shot" type, which usually incorporates a frangible portion that breaks upon valve operation (Figure 25), and the

quick-disconnect type, which is installed so that it will be

triggered (released) during the crash sequence (Figures 20 and

26). Some valves in use today have both these features incorporated into their design. Each specific fuel system design will dictate which of the two types of valves can be used. In

either case, the valves must be installed in a manner that pre- cludes inadvertent operation.

41 [ half)Vle. •raaa(tank O•. pakiTtink Tank fitting

Frangible section

Breakaway valv

(hose half)

Figure 25. "One-shot" self-sealing valve. (Load

on hose or lower valve body causes separation at frangible section.)

Release ring te*"as rn T rig ggeer ccaa blý e

S~Flex ."Quick-disconnect valve hose

Figure 26. Cable-actuated quick-disconnect valve. (If a

triggered cable system is used, its location

must be carefully selected to prevent inadvertent valve actuation during normal aircraft

operation and maintenance.)

67

1. *
Self-sealing breakaway valves should be located at each fuelcarrying tank outlet and at locati.ons within the fuel line network where extensive displacement is foreseeable, such as wiing

roots or engine compartments. The purpose of these valves is

to prevent rupture of the tank, hoses, or fitting comporients

by placing a "safety fuse" in the load path.

A self-sealing breakaway valve should be used to connect two

fuel cells in a direct side-by-side arrangement, especially if

there is a high probability that structure failure will occur

in the immediate area of the cells. Figure 27 shows a breakaway valve mounted in such a cell-.to-cell installation.

Rib or similar

structure

Fuel cell wall

Frangible section,

SValve body Valve body

Aircraft tank - Sump area

cavit" loor

Aircraft outer skin

Figure 27. Typical cell-to-cell self-'sealing

breakaway interconnect valve.

Tank-to-line interconnect valves should be recessed sufficiently into the tank so that the tank half is flush with the

tank wall or protrudes only a minimal distance beyond the tank

wall after separation. This feature reduces the tendency of

the valve to snag on adjacent structure during the crash sequence.

68
The frangible interconnecting member of each of these valves

should be sufficiently strong to meet all operational and

service loads of the aircraft within a reasonable margin but

should separate at 25 to 50 percent of the minimum failure

load for the weakest component in the fluid-carrying line.

Figure 28 illustrates a sample breakaway load calculation.

Airframe structure

Tank

Hose end

coupling

Metal tank fitting

Breakaway valve

Frangible section

i ITEM LOWEST FAILURE LOAD (LB)* FAILURE MODE

Flex hose 3000 Tensile breakage

Flex hose 1500 Pull out of end

fitting

Tank fitting 7500 Pull out of tank

Hose end coupling 1650 Break (bending) Breakaway valve 2500 Pull out of tank

fitting

Breakaway vdlve Not more. than Brea]. at fran1500 = 750 gible section

2

i Not less than

1500- -- =375

*Loads may or may not be representative; values are for

explandtory purposes only.

Figure 28. Typical method of breakaway load calculation

for fuel tank-to-line breakaway valve.

69 ...

..• • . . j ; P ./' - .. f ..A . . .
Each valve application should be analyzed to assure that the

probable separation load will be exerted in a direction and

manner to which the valve is best suited. These loads, whether

tension, shear, compression, or combinatiuxis thereof, are obtained by analyzing the aircraft for probable impact force and

direction and by determining the consequent structural deformation around the valve.

Care should be taken to select valves that allow only a minimum

amount of spillage upon separation. The valve should permit

no external leakage when partially separated. For this reason,

valves with a very short triggering stroke are superior to

those with a long stroke.

Operational pressures are dependent on specific applications,

but the valve designs can take advantage of the available line

pressure to assist in keeping the self-sealing mechanism closed.

As in all valve design, light weight and minimal pressure drop

are major design objectives, but the resistance of the valve

to direct impact or to high compressive loads should not be

sacrificed for the sake of weight reduction.

4.2.3.2 Ventp: Vent systems become involved in the crash fire

episode when-the aircraft rolls far enough to one side to allow

fuel to drain out of the systems and/or when the vent lines

fail.

Vent line failure often occurs at the point of exit from the

tank. Failure at this point can be reduced by using short,

high-strength fittings between the metal insert in the tank

and the vent line. The vent line should be made of wirecovered flexible hose and should be routed in such a manner

that it will not obviously become snagged in displacing structure and torn from the tank. Self-sealing breakaway valves

also can be placed at the tank-to-line attachment area. This

approach IIbecomes mandatory if there is danger of the tank being •, torn free of the supporting structure.

Vent lines should be routed inside the fuel tank in such a manner that, if rollover occurs, spillage cannot continue. This

can be accomplished with siphon breaks and/or U-shaped traps

in the line routing.

Recent developments indicate the feasibility of placing vent

valves inside the fuel tank. These valves are designed to operate in any attitude and to allow a free flow of air while

prohibiting the flow of fuel. They are particularly advantageous during rollover accidents, and can be used in lieu of

flexible lines, breakaway valves, and all other alternate considerations. One current type of vent valve is illustrated in

Pigure 29.

70

1

....... 1
V. ent

Frangible section

Tank fitting

.- Tank wall

Vent valve body

~.Floating ball

Side view iiii. !i:';':•

; :,eavy ball

Vent valve body

Bottom view Ball

---- Air passageway

Ball retention pins

Figure 29. Vent valve.

If the fuel system is to be pressure refueled, it should be

noted that a bypass system for tank overpressurization will

have to be used. This capability can be built into the vent

valve or can be incorporated in a separate unit. In either

case, however, care must be taken to ensure that spillage resulting from overpressurization due to tank compression during

a crash is released away from aircraft occupants and ignition

sources.

71

~~I'I
4.2.3.3 Boost Pumps: Fuel boost pumps fall into two general

categories, There are the tank- or line-mounted types, which

pressurize the fuel lines, and the line- or engine-mounted type,

which suck fuel from the tank and lines, creating a slight negative pressure in the fuel lines. Suction fuel systems pose a

much lower threat in regard to crash fires; however, both systems can pose potential problems. Some boost pumps in use today are installed In the fuel tank and are rigidly bolted to

the aircraft structure. Crash damage to the pump can cause

fuel spillage and also supply electrical sparks for ignition

of fuel.

The state of the art in fuel system design has shown that electrically driven boost pumps can be eliminated. Air-driven

boost pumps and engine-mounted suction-type boost pumps now in operation are much less hazardous alternative solutions.

If design requirements dictate that a boost pump be installed

in the fuel tank, it is suggested that the pump be air driven

and that it be rigidly bolted to the fuel cell only. If the

pump must be supported or attached to the aircraft structure,

a frangible attachment should be used, as shown in Figure 30.

4.2.3.4 Filler Necks: The filler necks commonly used on

present-day aircraft can, and frequently do, cause fuel tank

failure. Typical filler neck installations place the cap at

one end of the filler tube and the tank at the other end. During periods of structural displacement, the neck can be pulled and torn from the tank, leaving an opening in the tank wall.

To prevent fuel spillage, it is imperative that the filler cap

remain with the tank. To do so, it must be mounted at, or

slightly below, the tank wall surface.

Although the use of filler necks is not recommended, certain

aircraft configurations require their use. It is suggested that a frangible type be devised, as shown in Figure 31. Alternatively, a check valve can be placed in the tank filler

opening as shown in Figure 32. Another suggestion for filler

attachments is the frangible ring concapt presented in Figure 33.

4.2.3.5 Quantity Sensors: Accident investigations have shown

that quantity sensors cause two types of tank failures. The

first type of failure, which is common to most quantity sensor

installations, involves the rigid attachment between the sensor

entry into the tank and the aircraft structure. This rigid

coupling cannot accommodate much structural displacement without inducing a tearing failure in the fuel tank. It is necessary, therefore, that a frangible structure be used for this ti type of tank attachment (see Figure 34). An alternate approach

is to make the probe mounting attachment frangible.

72,"

........................... ......
Boost pump

Pumpto-arcraPump-to-tank Pup-o-irrft attachment Airframe

structure attachme bl srctr

(franqible)

Figure 30. Frangible attachment of air-driven

boost pump to airframe structure.

Aircraf t

skin

Filler neck

Access door

toprille capm

enrFuel tank

Figue Fangblefiler 3. eckinsalgbelati n.

A-j

L)n 

wal
Aircraft skin

Frangible filler neck

Failure plane Filler cap Floor i

Check valve

(trap door, spring- Fuel tank loaded closed)

Figure 32. Check valve installed in conjunction

with frangible filler neck.

Aircraft structure

Frangible .T

ring-to-A/C structure Tank Failure plane

bolt'------*- Nut ring

Cap retention

ring

Frangible ring

To

Frangible tank

ring-to-fuel LFailure

tank bolt points

To aircraft

Figure 33. Frangible ring attachment for

installing fuel tank fillers.

The second type of sensor-induced tank failure is the punctur- ing of the tank by the long, rigid, tubular sensing probes in

use in many aircraft. Corrective approaches to thi3 problem [ include mounting the probe at a less hazardous angle or using curved, frangible, low-flexural-rigidity probes, or probes

74

A64

. • 1• ' '; .• ,- . .
Quantity Fuel cell Ri-vtindicator ,

Nut ring Aircraft skin

Frangible ring

Frangible probe

... ...

Figure 34. Crashworthy fuel quantity sensor.

equipped with load-spreading shoes, fuel counters, and floatand-arm type sensors. While the new crash-resistant tanks have

greatly reduced this problem, it still poses a hazard that

should be remedied in the larger cells.

4.2.3.6 Sume Drains: Sump drains are a frequent source of

fuel spila ge because their design dictates that they be located at the lowest point in the tank, in close proximity to

the most probable impact area. Figure 35 illustrates some design concepts that permit maximum drainage without the drain

protruding beyond the face of the tank.

4.2.3.7 Fuel Strainers and Filters: In-line fuel strainers

should not be located in the enTg-- compartment if such a prac- tice can be avoided. Engines are sometimes torn loose during crash impact, and the strainers located in the compartment are

susceptible to damage from the displaced engine. Mounting of

the strainers directly on the engine is not desirable. The engine location might afford some protection during a crash, but

its proximity to the hot engine surfaces creates an additional

hazard from ballistic hits. Strainers should have a structural

attachment capable of withstanding a 30 G load applied in any direction to minimize the possibility of their being torn loose

during crash impact. Self-sealing breakaway couplings should

be used to attach fuel lines to the fuel strainers if there is

a probability of line damage at this point.

75

U KW
Standard drain cock protected by ramped fitting

Fitting

Drain cock

Fuel cell

Push-twist type, flush-mounted drain valve

Rotate to

lock open

tDepressor tool

Valve closed Valve open

' 'I

Figure 35. Design concepts for crashworthy fuel drain cocks.

4.2.3.8 Capsand Access Covers: These items play a major

role in crashworthy fuel containment. Since they function as

seals for tank openings, their failure could be catastrophic.

Caps having a minimum rating of 75 psi or greater should be

used. Access covers must not be the weak link in the fuel tank.

They must be capable of carrying loads equal to or greater than

those which the tank can withstand.

76

.e 3.k. 
4.2.4 Fuel System Full-Scale Crash Test

Consideration should be given to conducting a crash test with

the complete crashworthy fuel system in enough of the airframe

to create a realistic situation. Since the subject aircraft

can crash in a variety of attitudes and speeds, the attitude

and impact velocity for the fuel system test should be representative of the attitudes and velocities used in the crashworthy design of the overall aircraft. The recommended design velocity changes are listed in Table 5. The reader is referred

to Volume II for a complete discussion of crash design conditions.

TABLE 5. SUMMARY OF DESIGN VELOCITIES

FOR ROTARY- AND LIGHT FIXEDWING AIRCRAFT

Velocity

Impact change

Direction (ft/sec)

Longitudinal 50

Vertical 42

Lateral* 25

Lateral** 30

*Light fixed-wing.

**Rotary-wing.

4.3 OIL AND HYDRAULIC FLUID CONTAINM4ENT

Oii and hydraulic fluid spillage often occurs in aircraft accidents. Fortunately, these fluids are carried in much smaller

quantities than fuel. However, they are easily ignited; oil

is usually carried hot, which makes ignition easier; they are

pressurized in places, which converts them into mists when

they are released, making ignition easier; and they are often

carried near the hot engine, which can readily provide ignition.

SWhen oil or hydraulic fluids are ignited, they, by themselves,

constitute a low threat to aircraft occupants. But, unfortunately, they function as ignition sources for other combustibles, especially spilled fuel. Further, they migrate throughout the wreckage, carrying with them flames that otherwise

would not be present. Oil and hydraulic fluid spillage, therefore, should be prevented at all reasonable cost.

77 A

* 4 - -:/ -'" ' •, •,".' l "'
The crashworthy design criteria presented for the fuel tanks,

lines, and supportive components all apply to these fluid systems, with one possible exception. Because of their relatively small capacities, properly protected metal tanks may be used.

It should be recognized, however, that metal tanks are punc- tured easily and are not tear resistant. If tank puncture is

likely, several alternatives are available: a crashworthy

tank, like the fuel tank, can be used; the tank can be relocated to a safer area; or the tank can be shielded.

Experiments have been performed to determine the practicality

of shielding a metal oil tank with a 1/2-inch-thick felt cover

made of ballistic nylon, as shown in Figure 36 (Reference 18).

As an added degree of spillage protection, the outside surface

of the felt was coated with a thin layer of polyurethane resin

to make it leakproof. Since preliminary experiments proved satisfactory, a similar system was crash-tested in a U. S. Army UH-l helicopter. The tank sustained severe impact damage, rupturing a tank seam. The spillage leaked out into the felt

cover, but did not escape from the felt due to the polyurethane

coating. This system is simple, light in weight, easy to install, and relatively low in cost. A similar felt tank cover

is now being used with a high degree of success on all Indianapolis-type racing cars.

When metal lines must be used in these systems, they should be

designed to incorporate a coil or two of extra line length so

that the line can stretch to accommodate some structural distortion. Also, the lines should be attached to the airframe

with clamps that will fail and release the fluid lines before

the line itself fails, thereby allowing the line to change its

routing to help accommodate structural distortion.

Hydraulic fluids that inherently resist burning should be used

whenever possible (Reference 19). Most of those fluids, however, have operational and maintenance problems associated with

their use. Therefore, designers may wish to consider the

trade-off of using conventional hydraulic fluids, as compared

with using fire-resistant fluids. It should be noted that,

even though the new fluids are fire resistant, most of them

18. Robertson, S. H., DEVELOPMENT OF A CRASH-RESISTANT FLAMMABLE FLUIDS SYSTEM FOR THE UH-lA HELICOPTER, Dynamic

Sciencel USAAVLABS Technical Report 68-82, U. S. Army

Aviation Materiel Laboratories, Fort Eustis, Virginia,

January 1969, AD 688165.

19. Military Specification, MIL-H.-83282, HYDRAULIC FLUID,

FIRE RESISTANT SYNTHETIC HYDROCARBON BASE, AIRCRAFT, Department of Defense, Washington, D. C., 22 February 1974.

76

"".'

.•,I 

I:
Filler cap I

Liquid-tight, ... .. .

fire-resistant coating: N

Felt.

oil tank Ri

NOTEM The felt is not bonded to the oil tank.

Figure 36. Felt oil tank cover.

will still burn, especially when in a mist state. The characteristics of each fluid must be studied before the final tradeoff decision is made.

4.4 FUEL MODIFICATION

One method for decreasing the postcrash fire potential is to

decrease the susceptibility of aircraft fuels to dispersion

and atomization, reducing the formation of combustible fuel/air

mixtures. This can be done through the use of fuel modification additives. These modifying agents have been classified

as either antimist, emulsification, or gelling additives. The

blending of these additives into standard aviation fuels pro- vides fuel properties that decrease the tendency to disperse,

atomize, and form fuel mists following crash-induced fuel system failures. As a result, retardation of fuel mist fireballing and fire propagation can be achieved.

II 79
Several studies have shown the feasibility of providing postcrash fire protection through the use of antimist fuels, emulsified fuels, and gelled fuels (References 20, 21, 22, and 23).

This approach has been fairly successful when used with low

volatility fuels such as JP-5, JP-8, and Jet A. However, modification of highly volatile fuels, such as JP-4 and aviation

gasoline, has not been effective. Emulsified and gelled fuels

have received little attention recently due to their inherent

system compatibility problems. Further, consideration of modified fuels has declined since the development and use of

crashworthy fuel systems in rotary-wing aircraft. However,

the possible use of antimist fuels in fixed-wing aircraft where

crashworthy fuel tanks are less feasible has generated recent interest.

Although turbine engine performance is not adversely affected

by use of antimist fuel blends, it has been found that these

fuels must be degraded before starting and restarting a turbine

engine with a standard fuel system. During startup, the characteristics of the antimist fuel suppress the atomization of

the fuel through the fuel nozzle, thus starving the initial

ignition. To alleviate these problems, processes to reverse

the antimist fuel blend so that the blended fuel can be brought

back to the neat state prior to introduction into the aircraft

fuel feed system are being investigated.

20. Weatherford, W. D., Jr., and Wright, B. R., STATUS OF RESEARCH OF ANTIMIST AIRCRAFT TURBINE ENGINE FUELS IN THE

UNITED STATES, in Aircraft Fire Safety, AGARD Conference

Proceedings No. 166, North Atlantic Treaty Organization,

Advisory Group for Aerospace Research and Development,

Neuilly-sur-Seine, France, October 1975, pp. 2.1-2.12,

AD A018180.

21. San Miguel, A., ANTIMISTING FUEL KINEMATICS RELATED TO

AIRCRAFT CRASH LANDINGS, Journal of Aircraft, Vol. 15,

No. 3, March 1978, pp. 137-142.

22. Shaw, L. M., SAFETY EVALUATION OF EMULSIFIED FUELS, Dynamic Science; USAAMRDL Technical Report 71-29, Eustis

Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, June 1971,

AD 729330.

23. Shaw, L. M., SAFETY EVALUATION OF ANTIMIST FUELS, Dynamic

Sciencel Report 9130-73-112, U. S. Army Mobility Equipment

Research and Development Center, Aberdeen Proving Ground,

Maryland, November 1973, AD 773035.

80

• .. ... ., ... : '••a `• • ['•` * .•.AVI ,r . • ; e!' ,, -• , i L. :M.7,h' ' - .
4.5 IGNITION SOURCE CONTROL

Flammable fluids will ignite throughout a wide range of temperature, pressure, atmospheric composition, and ignition source

conditions. Generally, ignition of spilled combustibles during

the crash occurs from one or more of the following: electrical

sources, flames, hot surfaces, and friction sparks. Components

usually involved in the ignition process include the engine, exhaust system, heater, battery, wiring system, and various

light bulbs.

4.5.1 Electrical Sources

The aircraft electrical system is a potential crash-fire igni- 2 tion source because it is distributed extensively throughout the aircraft and because electrical discharges are able to concentrate a high amount of energy into a small volume.

Disruption of a current-carrying electrical circuit can result

in fuel ignition by electrical sparks and arcs that are released when exposed wires contact grounded surfaces. Ignition

also can be provided by wires that have been heated either by

short circuiting or by normal means, as in an incandescent

light filarment. The common incandescent filament in a landing

light is hot enough to ignite fuel 0.75 to 1.50 sec after bulb

breakage.

Perhaps the most important aspect of an electrical discharge

ignition source is the great amount of energy present compared

with the small amount actually required to produce fire ignition under Adeal conditions. Approximately 0.15 millijoule

(0.11 x 10 ft-lb) is the minimum energy for spark ignition "under ideal temperature, pressure, and mixture conditions.

The ignition potential of the aircraft's electrical system may

be reduced by aircraft modification at the system level and at

the component level. The system level approach is concerned

with de-energizing electrical generation or storage systems,

whereas the component level approach is concerned with component location and environment.

4.5.1.1 System Level Approach: Reduction of crash-fire ignition by the electrical system can be achieved by removing from

the electrical circuit all electrical generation or storage

systems before or during the early phases of the crash sequence.

The dc-energizing can be accomplished by opening the electrical

circuit at the output terminals of each energy-producing component.

-------
The time required for this de-energizing operation is of utmost

importance. Crash-fire data previously reported by NACA, using

both aviation-grade gasoline and low-volatility fuel, indicate

a minimum time of 0.7 sec between impact and fire ignition with

the electrical system as the source. During helicopter crash

tests by AvSER, it was observed '-hat fire started to propagate

approximately 0.58 sec after ground impact. During tests with

simulated fuels, macsive fuel spillage was in progress as early

as 0.20 sec after impact. Therefore, each de-energizing device

must be capable of activation within a maximum time of 0.20 sec.

,! The primary items to be considered for de-energizing are the

batteries, generators, and inverters. Several precautions must

be taken. Since the battery can remain a potential ignition

source for hours after a crash, ends of wires severed from batteries must be prevented from contacting the structure and

thereby providing a new igni.tion source. The generators and

inverters cannot be satisfactorily de-energized by simply opening field circuits. There is a considerable time lag (0.385

sec for a rotating inverter) between DC input cutoff and AC

output termination (Reference 24). Therefore, for complete

safety, these components must be disconnected from buses on their output sildes. NACA also recommended that consideration

be given to grounding the armatures of main electrical components close to those components (Reference 25).

Magnetos and igniters are of special interest, since they are

high-energy sources of ignition. If these componcnts were deenergized, the fuel in the engine during the crash event would

not be ignited. However, raw fuel then would be pumped into the hot exhaust manifold, resulting in a fire. Recent crash- fire research has demonstrated that it is better to turn the

fuel off and to leave the ignition system on throughout the

crash sequence (Reference 24).

Relays can be used to de-energize components and to activate

other inerting elements. In the case of batteries, only nonessential buses should be disconnected initially. Power must

24. Robertson, S. H., et al., THEORY, DEVELOPMENT, AND TEST

OF A CRASH-FIRE INERTING SYSTEM FOR RECIPROCATING ENGINE

HELICOPTERS, Aviation Safety Enginaering and Research of

Flight Safety Foundation; TRECOM Technical Report 63-49,

U. S. Army Transportation Research Command, Fort Eustis,

Virginia, December 1963.

25. Pinkul, I. I., et al., ORIGIN AND PREVENTION OF CRASH

FIRES IN TURBOJET AIRCRAFT, NACA Report 1019, Lewis Flight

Propulsion Laboi.atory, National Advisory Committee for

Aeronautics, Clevelaaid, Ohio, 1958.

* .82
A ..

be provided to other elements of the crash-fire prevention system until these elements have completed their design functions.

A time-delay unit can be ubed to cut off power to inerting elements and to ground the disconnected buses. An alternative

to a relay contact is the explosive cable cutter shown in Figure 37. The electrical system inerters must, in any case, be

capable of resetting components in the event of inadvertent operation.

To system

Ground -ii

,-Electrical de-energizer

Actuation

switch Explosive cable

cutter

From battery U -w.'o system

Reset switch

manual reset

Detail A

Figure 37. Cable cutter de-energizing method.

A 4.5.1.2 Component Level Approach: The ignition hazard associated with the electrical system can be reduced at the component level by controlling component location and environment.

The following guidelines are applicable to batteries, inverters, generators, alternators, magnetos, igniters, radar, antennas, and lights.

Components should be located above and away from flammable

fluid sources. Leaking flammable fluid should not come in contact with electrical equipment or wiring as a result of gravity, airflow, or battle damage. The electrical system components should be located, and suitably mounted, in areas where

anticipated impacts will be minimal and where maximum anticipated structural deformation will not result in structural

83

}iI

-T ... '- ... " .' . -, . . ...... . ... .. . .. . . . ..... ,. • _.. .,. .. . ... .'.•,• ,."7]
impingement on either components or wiring. Wires should have

6-in.-diameter loops near their component connections to accommodate any wire tensioning resulting from structural deformation. All wire connections should be made on a component's

least vulnerable side. Batteries, inverters, and generators

should be mounted in compartments lined with tough, nonconductive shields. The shields will prevent sparking between terminals or severed wires and the aircraft structure. The components should be mounted to the aircraft with structural

attachments capable of withstanding 30 G loads in any direction.

Electrical wires should be routed along the strongest structural members and should not, in general, traverse areas of

anticipated severe structural deformation, e.g., in leading edges of wings or in the lower regions of the fuselage. Wires

that must pass through areas of anticipated structural deformatiorn should be approximately 20 to 30 percent longer than necassary. The extra length should be accumulated in the form of

loops or S-shaped patterns and located at the areas of anticipated structural deformation. When wires pass through structural openings or bulkhead holes, the openings should be 8 to

12 times larger than the wire diameter and appropriate grommets

0 should be provided. The wires should be attached to the aircraft structure with clamps or ties that will fail before

breaking the wire. Nonconductive shields should surround all

areas where wire abrading or cutting may occur. Wires should

not be routed near flammable fluid sources.

The mounts for antennas and lights should be attached to the

aircraft with frangible structures. The wires should incorporate a shielded covering and/or a breakaway capability. A suggested installation technique for a rotating beacon is illustrated in Figure 38. Structural impingement upon the component

will be difficrult because the frangible mounting structure will

allow the beacon to displace. The extra wire contained in the

loop can allow for considerable beacon movement without failing;

if massive displacement is anticipated, shielded failure points

can be used. Theoe same techniques apply for all similar types

of components.

4.5.2 Engine

The two principal engine ignition sourcer are (1) intake, combustor, and exhaust flames, and (2) hot ital surfaces. The

differences between these two relate to the time that these

sources persist after a crash, the manner in which ignition

occurs, and the mode of propagation of the resulting fire out

of the engine.

84

4 ... ... 

.-
Rotating beacon

Frangible clip %r

2Frangible structure

6-in. loop of 

extra wire • . •

Electric wiring

Flexible nonconducStive shield Wire and insulation

Wire not \Failure point

shield Wire bonded 4 .,.~ *..to shield

Electrical wiring LFailure point\-Wire and insulation

Hot side

"` Figure 38. Rotating beacon installation.

4.5.2.1 Flames: Engine inlet and exhaust flames are responsible for the Ignition of many crash fires. During the crash

sequence, flames often appear at these locations due to engine

breakup or rapid changes in engine loading, as can occur when

"a drive shaft is severed or a propeller is sheared.

iAll

85

* -.
Flames also appear at these locations when engines ingest

spilled fuel. Turbine engines are highly susceptible to fuel

ingestion because of the relatively long period of time required for the turbine to coast to a stop. The ingested fuelair mixture enters the downstream end of the combustor, where

flames may persist for up to 18 sec after fuel cutoff. The ingested mixture then burns in the tailpipe downstream of the

turbine (Reference 26). Also, under the proper conditions, the

combustor flame may propagate upstream through the ingested mixture and exit at the engine inlet.

The occurrence of engine inlet and exhaust flames and the resulting ignition hazard can be reduced by stopping the fuel

flow, by inerting the flame source, and by providing shielding

to prevent fuel spillage from entering anticipated flame areas.

The engine fuel valves should be closed, but the ignition system should be left on to permit normal burning of ingested fuel

in order to prevent undesirable exhaust or inlet flames.

4.5.2.2 Hot Surfaces: The probability of flammable fluid

ignition due to contact with a heated surface during a crash

is high and can remain so for several minutes after a crash.

The circumstances leading to ignition are somewhat involved;

generally, they are dependent upon the type of flammable fluid

involved, temperature of the fluid, composition of the heated

surface, temperature of the heated surface, geometry of the

heated surface, ratio of the fuel to air, and the degree of

fuel atomization.

Ignition temperatures vary widely. As a general rule, hydraulic and lubricating oils ignite at lower flat-plate temperatures than aviation gasoline. JP-4 also has a lower flat-plate

ignition temperature than aviation gasoline. The lower grades

T. of gasoline have lower ignition temperatures than the higher

grades. Kerosene has a lower ignition temperature than JP-4.

V, The ignition temperature of a flammable fluid is directly related to the initial fluid temperature. While the fuel temper- ature can vary considerably, depending on temperature at altitude, on the ramp, and at the storage facility, the temperature

of the oils is of more concern. As mentioned, oils can ignite

at a temperature lower than most fuels, and since they are carried in the heated state, low hot-surface temperatures and exposure times will provide ignition. Oil fires can, in turn,

act as ignition sources for the fuel.

26. Black, D. 0., CRASH-FIRE PROTECTION SYSTEM FOR A J57 TURBOJET ENGINE USING WATER AS A COOLING AND INERTING AGENT,

NASA Technical Note D-274, National Aeronautics and Space

Administration, Washington, D. C., February 1969.

86
The time between fuel contact with the heated surface and fuel

ignition is directly related to the temperature of the heated

Ssurface. As shown in Figure 39, the hotter the surface, the

faster ignition can occur. Also, it can be seen that ignition

can occur at a much lower surface temperature if the exposure,

or residence time, is longer.

1500 --

1400

1300 " 0 JP-4 Fuel

(unpublished data)

1200 -.. -Extrapolated

4J 1100 - ,

*• 1000 -

0• goo . ,,!!11 ..

S.• ,~~oo,,ul

*9b::JP-4 fuel -

0 7 0 0 ; . . A T,• """ -! '!- 

S600 • Pressure,

500 - --

400 --

t .300 - - .01 .1 1 10 100

Minimum time delay before ignition, sec

Figure 39. JP-4 ignition delay versus surface temperature.

(Taken from Reference 25)

87 A

S.. . .. . .. . :.t .,l ..•l• • i.V ... . ;. -,• .. . ..

- • ,,.••,. w " •,. J • 417'•• 
The ratio of fuel to air also governs the probability of ignition. The designer must assume that the proper ratio does

exist somewhere within the spillage area.

The potential hot surface ignition sources on aircraft with

reciprocating engines are the intake systems, the exhaust gas

disposal systems, heaters, and the higher temperature regions

of the cylinders. Ignition sources on turbojet engines include

the internal areas downstream of the compressor, and external

areas from the compressor aft, including the tailcone and tailpipe, and, in some designs, the bleed air system. The gas

flow through a turbojet engine may be too rapid to permit the

ignition of ingested combustibles on hot metal in contact with

the main gas stream. However, a portion of the engine airflow

is diverted to hot surfaces not in the main gas stream where

ignition may occur.

The hot surface ignition hazard can be reduced by methods analogous to those used with the inlet and exhaust flame hazards.

An inerting system can be used to reduce the temperatures of

hot surfaces to predetermined acceptable levels and to surround

the hot surfaces with an inert atmosphere to prevent ignition

from occurring should flammable fluids be spilled on these surfaces. In previous studies, hot surfaces have been cooled to

temperatures ranging from 400*F to 7600F with satisfactory results (References 24 and 27). The temperatures to which hot

surfaces must be cooled to prevent ignition must be determined

as a function of the fuel and the engine configuration to be

used; however, 400UF should be used as the upper limit for safe

hot surface temperatures.

Shielding also can be used to prevent spilled flammable fluids

from reaching the hot surfaces.

4.5.2.3 Inerting Systems: The function of inerting systems is

to render ignition sources harmless and, therefore, to prevent fire or explosion. Inerting systems can be designed to surround hot surfaces with an inert atmosphere. With an inerting

system, there is not sufficient oxygen to support combustion

when flammable fluids contact the hot surfaces. These systems

also can be designed to perform the additional function of cooling the hot surfaces to temperatures below the ignition temperII *1 atures of flammable fluids. Knowledge of temperature gradients

and cooling rate characteristics for each particular engine design is required in order to design an adequate cooling and inerting system.

27. Pinkel, I. I., et al., MECHANISM OF START AND DEVELOPMEN4T K OF AIRCRAFT CRASH FIRES, NACA Technical Note 2996, Lewis

Flight Propulsion Laboratory, National Advisory Committee

for Aeronautics, Cleveland, Ohio, 1953. 

I

88

I.
Ii

The system must be capable of providing a high-discharge-rate liquid spray for rapid cooling, a lower-discharge-rate liquid

spray over the more massive engine surfaces for a longer period

of cooling and inerting, and a follow-up inerting spray. Water

is a preferred coolant because of its high latent heat of vaporization, its availability, and its low cost. Additives to

the water can be used to protect against freezing and corrosion

of the piping system. Nitrogen and carbon dioxide are among

other inerting agents that have been used successfully as

follow-up sprays.

Several successful hot surface and/or flame inerting systems

have been designed, tested, and incorporated into current military aircraft. A schematic diagram of a hot-surface inerting

system used on a reciprocating engine is shown in Figure 40

(Reference 24). Systems such as this have been installed on

aircraft and crash tested. They successfully inerted the en- V

gine and exhaust systems, thus preventing crash fires.

Testing of a pyrotechnic gas-generator-type extinguisher system

indicates that it offers performance improvements over the

pressurized nitrogen-type system (Reference 28). The pyrotechnic syctem was more effective at low temperatures and with less

volatile extinguishing agents, and it eliminated problems associated with the mixing of thednitrogen and the liquid agent.

Flame-source inerting systems are designed to extinguish combuster flames, which linger long after the fuel has been cut

off. The flames are a result of the ignition of fuel that remains in the fuel manifold and continues to drip into the com- "bustor. A schematic diagram of a flame-source inerting system

used on a reciprocating engine is shown in Figure 41. Upon

actuation by either a manual or a crash-actuated switch, highpressure CO , or a comparable inert gas, is released into the

engine air Zntake. Concurrently, the high-pressure gas is used

to activate linkages that close the fuel inlet valve, the oil

inlet valve, and the air intake opening. The large volume of

inert gas released into the cnginQ interior quickly dilutes

the incoming air to the point of nonflammability, thereby eliminating the flaraes. As the engine continues to coast to a stop,

the inert mixture is pumped through the engine and expelled at

the exhaust outlet.

28. Klueg, E. P., et al., AN INVESTIGATION OF IN-FLIGHT FIRE

PROTECTION WITH A TURBOFAN POWERPLANT INSTALLATION, Report

No. NA-69-26, Department of Transportation, Federal Aviation Administration, National Aviation Facilities Experimental Center, Atlantic City, New Jersey, April 1969,

AD 686045.

89

L ii .... .. ... .. i
N high-pressure • ~re servoir

Regulators

N2 high-pressure

H2 0 accumulator manifold 2 reservoir

N2 low-pressare

manifold

N 2 pressure transducer

Filter

Explosivoperated

"valves

N /H 0

aspiratorBle

Bleed valve Bleed valve

N2 /H 2 0 aspirator

H2 0 manifold panifold 11 2 spray manitold i 2 

ii•'" -To spray

distribution

manifold, 12

per side

IHO H20 MMEZ3r ' - : i

N2 high pressure

N2 low pressure [

it.,,. I; / 2 •a

Figure 40. Hot surface inerting system.

90

I,;

PII - L 
CO2 pressure reservoir

Explosiveoperated vl

Piston in fired position

Distribution Oil

manifold cutoff

Fuel cutoff valve vavveae _i

Carburetor

closinq lever Itra

" blower

S~Engine plenum area.

Figure 41. Flame-source inerting system.

In a turbojet application, the combustor flames were eliminated by providing for rapid fuel shutoff and draining of the

fuel manifold. The system is shown in Figure 42 (Reference 26).

The fuel was shut off by a pneumatically operated valve installed in the fuel line between the engine fuel-control unit

and a modified pressurizing and dump valve. Simultaneously,

the manifold drain valves and the modified pressurizing and

dump valve opened and vented the fuel manifold overboard. The

91
'.41

.5 

.1 > Q) 

4-J fO . J

0 r-I00 0

4-I 0 " -

.1Jp >I44

44 0

000

-4- cýI 44 

r- (

92
combustor air pressure, available at the instant the fuel shutoff valve closed, then reversed the fuel flow in the nozzles

and manifold through the overboard fuel drains. The combustor

flame was extinguished in 0.23 sec by this method.

As is the case with electrical system de-energizing devices,

the inerting systems should be operable within 0.20 sec of the

sensing of a crash.

4.5.2.4 Shielding: Shielding is an effective method of preventing flammable fluids from reaching potential ignition

sources. Shielding can take many forms; however, there are

three general methods in use, with a fourth now in the development stage.

The first method of shielding uses baffles. Metal or other

rigid paneling will not satisfy the shielding requirement because of its inability to maintain an effective seal in areas

of large structural displacement. Sealed curtains or baffles

made of fire-resistant cloth or similar material can perform satisfactorily. The only requirement is that they must seal

all openings through which flammable fluids could travel to an

ignition source. To accommodate the anticipated structural

displacement, it is suggested that all curtains and shields be

at least 30 to 40 percent larger than the minimum size required

to protect a given area. Figure 43 illustrates how the flexi- ble curtain concept was used to keep fuel from entering the

occupiable areas on an experimental test helicopter.

The second method of shielding uses spillage flow diverters or

drip fences. Once liquid has settled onto a sloping surface,

it flows to the lowest point. It can flow on top of a surface,

or it can cling to the underside. In either case, it can

travel a considerable distance to an ignition source. Chordwise drip fences should be located on the wing on each side of

wing-mounted engines. Drainage holes should be strategically

located within the aircraft structure to drain internal spill- 4 age. All areas containing electrical components should be surrounded with a spillage gutter, or drainage trough, to prevent

flowing spillage from entering those areas.

Each engine and exhaust system mount should incorporate a drip

fence. Figures 44, 45, and 46 illustrate several types of drip

fence fuel-flow diverters.

The third method of shielding uses nonconductive flexible paneling. This type of paneling should be used as a liner for

electrical compartments and other regions where electrical components are installed. It should surround areas of electrical

wire groupings such as terminal strips and power control areas.

Nonconductive flexible shielding also can be used for shrouding

93

" ... ... , ,, I .
Rivet attachment

Extra folds Fir' uti .

.... .. ...

Fuel tank area

Fuselage section

Extra material folds

Figure 43. Fuel baffle and fire curtain concept.

94
Figure 44. External drip fence.

Wing ski

Liquid-tight sealA

S iga r 45rnenl rpfne

ea5
Firewall attachment points

Drip fences

Typical

engine 4 mount

Figure 46. Fuel flow diverters.,

or enveloping electrical wiring. The shielding Bhould'be used

in all places where structural shift or collapse could cause

an impingement on electrical wiring or related components.

The fourth method of shielding uses protective coatings or

surfaces. Recent studies have produced materials that, when

heated or exposed to other environments, expand to insulate and

protect the surface to which they have been applied. Intumescent paints are an example of this form of shielding. This

field of protective surfacing is fairly new and is an area recommended for future research.

96
4.5,3 Heaters

Heating units often are provided in aircraft cockpits and passenger compartments. These units, which also supply deicing

air, may be either combustion or engine bleed-air types.

Bleed-air heaters normally use air from the compressor section

of the engine. Hot-surface ignition sources on turbojet engines are downstream of the compressor section, and if the temperature is below 4000F, the piping system that carries the

bleed air to a mixing chamber should not be an ignition source.

If a temperature survey indicates that the system produces temperatures above 4000F, suitable inerting and/or shielding

should be provided.

Combustion heaters will produce hot metal surfaces that should

be treated as potential ignition sources. The surfaces of the

heater that become hot enough during normal operation to cause 4

ignition of crash-released flammable fluids must be determined,

and a cooling and inerting system must be designed. The coolant must not be an irritant to aircraft occupants. A waterdetergent solution was used in the cooling and inerting system

described in Reference 29.

4.5.4 Sparks

Two types of sparks should be considered potential ignition

sources: friction sparks and electrostatic sparks. The friction spark is a particle abraded from a parent meterial through

contact with a moving surface. Initially, the particle is

heated by friction. If the friction is great enough, the particle can burn, thus increasing its temperature. Electrostatic

sparks result from the discharge of an electrostatic charge

accumulated on parts during normal operation. The discharge

is triggered during the crash when the parts are separated due

to crash forces.

i• • 4.5.4.1 Friction Sparks: Friction sparks become possible ig- nition sources when portions of aircraft structure are scraped

* along the ground. While all common metals can be abraded, not

all spark sufficiently to ignite spilled fluids. Ignition occurrence depends on the thermal energy of the spark. The thermal energy is a function of the bearing pressure with which

the metal is abraded, the slide speed of the metal structure,

29. Jones, R. B., at al., AN ENGINEERING STUDY OF AIRCRAFT

CRASH FIRE PREVENTION, Walter Kidde and Company, Inc.;

Technical Report 57370, Wright Aeronautical Development

Center, Ohio, June 1958, AD 155846.

97

. 4 x 
the hardness of the metal, and the temperatures at which the

metal particles will burn.

NACA has conducted research on the friction spark ignition hazard relative to crashed aircraft (References 27 and 30). Some

results of this research are listed in Table 6. These studies

indicated that aluminum was the safest of the metals tested,

since it produced no visible sparks and did not ignite combustible mists at the highest bearing pressure and greatest slide

speeds tested. Of all the metals tested, titanium ignited the

combustible mist most readily; however, stainless steel, chromemolybdenum steel, and magnesium all ignited the mist at slide

speeds and bearing pressures less than those expected during a

crash.

TABLE 6. MINIMUM CONDITIONS UNDER WHICH CERTAIN

ABRADED METAL PARTICLES WILL IGNITE

Minimum

bearing pressure Drag speed

Metal (lb/in.a) (mph)

Titanium 21-23 Less than 5

Chrome-molybdenum steel 30 10

Magnesium 37 10-20

stainless steel 50 20

Aluminum 1455* 40

*Ignition was not obtained with aluminum.

No data are currently available on the spark hazards of the

new composite materials. Tests such as those mentioned above

should be conducted with these materials to determine their

friction spark ignition potentials.

30. Campbell, J. A., APPRAISAL OF THE HAZARDS OF FRICTION.

SPARK IGNITION OF AIRCRAFT CRASH FIRES, NACA Technical

Note 4024, Lewis Flight Propulsion Laboratory, National

Advisory Committee for Aerunautics, Cleveland, Ohio, May

1957.

98

.. . .. ..
There are two practical methods of reducing the friction spark

hazard. One is to use shielding to prevent the fuel from

reaching the spark-producing area, and the other is to build

the probable contacting surface out of materials having little

or no spark-producing tendencies.

As stated above, aluminum was the least likely metal to ignite

spilled flammable fluids. Building all aircraft structures

likely to come in sliding contact with the ground out of aluminum can reduce the spark hazard; however, it must be pointed

out that aluminum also is easily abraded. It can tear when

sliding, thereby exposing other metals that might spark and

ignite the spilled fuel. Therefore, the areas most apt to come

in sliding contact with the ground should be reinforced so that

longer contact times are possible without skin failure due to

abrasion. Particular attention should be given to attachment

points for hoists, landing gears, and other components located

in anticipated impact areas. Also, particular attention should

be given to the location of steel bolts, nuts, and washers.

All too often an otherwise spark-free area is contaminated by

locating a spark-producing bolt or nut within it.

4.5.4.2 Electrostatic Sparks: During the course of the NACA

research, it was noted that electrostatic discharge from a

wheel strut caused ignition of a fuel mist and, ultimately,

the destruction of the test aircraft (Reference 27). This ignition source was produced by a combination of environmental

conditions that would occur infrequently. It may be possible

to reduce electrostatic charge buildup by applying coatings to those parts of the aircraft likely to be separated in a crash.

Additional research is required to develop methods of eliminating this hazard.

4.5.5 Initiating Systems

A crash-fire prevention system should include an initiating system that senses the existence of crash-fire conditions and

causes action to be taken to suppress the ignition sources.

The initiating system must meet the following design requirements:

9 The system mu.zt not be capable of accidental operation as the result of malfunctioning sensors or short

circuits.

* The system must be designed to operate automatically

upon receipt of coincident signals from redundant

sensors.

99

[All,
"* The pilot must be capable of operating the system

manually and of overriding the automatic signals.

" The system must be designed for positive airborne

and ground check-out, with reset capability provided.

Sensors and the discriminating circuitry used to derive the

automatic signals must be carefully selected and developed.

The skill and knowledge of the designers also are important in

determining the location and installation of the sensors. A

discussion of sensors and criteria for aircraft application

are contained in Chapter 8 of this volume (Crash Locator Beacons).

The activating circuitry must be designed to avoid inadvertent

operation of the crash-fire protection system. Crash signal

redundancy is the key element in any such fail-safe system.

This design philosophy is illustrated by the activating circuitry shown schematically in Figure 47. This circuitry was

developed for reciprocating, multiengined aircraft by NACA

(Reference 31). A signal from any one of three switches will

result in the inerting of one of the engines. A fuel tank penetration switch indicates when the wing has been penetrated and

will result in de-energizing the electrical circuits within the

wing. Either the inerting of an engine or the de-energizing of

a wing's electrical circuits will cause a signal to be sent to

an arming control box. This signal must be combined with signals from two ground contact switches to actuate the entire inerting system. This requirement for simultaneous signals from

different types of initiating switches reduces the possibility

of the entire inerting system operating while the aircraft is

still in the air.

A schematic of activating circuitry that could be applied to

rotary- and fixed-wing single-engine aircraft is shown in Figure 48 (Reference 32). The average reading of four proximity

switches is compared with the aircraft's normal landing height.

31. Moser, J. D., and Black, D. 0., PROPOSED INITIATING SYSTEM FOR CRASH-FIRE PREVENTION SYSTEMS, NACA Technical Note

3774, National Advisory Committee for Aeronautics, Cleveland, Ohio, December 1956.

32. Drummond, J. K., STUDY TO DETERMINE THE APPLICATION OF

AIRCRAFT IGNITION-SOURCE CONTROL SYSTEMS TO FUTURE ARMY

AIRCRAFT, Dynamic Science; USAAMRDL Technical Report

71-35, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia,

June 1971, AD 729870.

100

........................................... ,. .,' c••G•J"'•• :' .. "
0

H- 4- W4 0 (D 4-4 

04 -*4 w0 0 4. J 4 4

p 4J d 0 4

0: A- U 0~ 4 . 44 H4 *,0r .

4.4 0 )Cf -) ( 44 4 4 4-)~~C)~4. 

4- 0 HJ .d U ;

o (a - o -

LA~~. -P-H0c

HH

u 0.u 0)

.94

Lo0

tr 0 0

101r

-- 4-J

0. -.- (3
I.Proximity 1 •-

IProximit~y 2

IProximity 3

SProximity 4

IAver aged•

Normal landing,

[height or lessI na

I Arm

Minimum tflying automatic

acceptable Not.ar~- 'u offe

durationlatns

S~ Hazard

0Iu ff•n

IPriot'y oeseni

input system eetia

"O"citchit

Figure 48. Potential activating circuitry

for single-engine aircraft.

If the average is less than the normal landing height for a

?!••,period of time that exceeds a preset minimum duration, an arm-

"I!,. ~ing signal is initiated. A second, independent arming signal " If ~ provided by a hazard switch is required before automatic opera- ,

i tion of the ignition-source suppression system. This hazard

switch may be any of the sensors previously discussed. Provisions also are included for pilot input to the arming signal

and for pilot override of the entire system.

102 •

.. . . ......... .. .. .. .. .. , .• tg • • i• ,• . .• .• ... I
5. INTERIOR MATERIALS

5.1 INTRODUCTION

While every effort should be made to prevent a major postcrash

fire by containing the fuel, as much protection as possible should be provided for the occupants in case a fire does start

at any time. Careful selection of interior materials can slow

the spread of smaller fires and give occupants time to evacuate

the aircraft safely or to be rescued by other personnel. The

protection afforded against in-flight fires is as important as

postcrash fire protection. Often, fire-hardening of the aircraft interior can result in a controllable in-flight fire incident rather than a catastrophic fire accident.

It would be desirable to present the designer with a concise

list of materials that should be used in aircraft interiors.

Unfortunately, this is not possible at the present time for

two major reasons. First, the selection of interior materials

is dependent on several varied and sometimes conflicting, design criteria. For instance, seat cushion materials must possess compressive modulus and rebound characteristics necessary for crashworthiness, restraint webbing must meet definite elongation criteria, and seat upholstery must possess a minimum

wear resistance. At the same time, these materials should pro- vide maximum fire resistance. Many materials currently available cannot meet all of the criteria simultaneously; thus, pri- orities must be established and trade-offs must be made. One

factor compounding this problem is that, at the time this volume is being written, there is no one place where all material

properties, including flammability data, are available. This

situation should be rectified shortly when the Transportation

r Systems Center data bank is fully operational (see Section 5.5).

The second major reason that precludes a listing of recommended

materials is that a great deal of activity has been directed

toward the development of materials and testing methods in the

last few years. This field is still very active and new materials and tests are being developed constantly. The designer

should be aware of this and select the best possible materials

for the aircraft interior. Many improved materials only now

are becoming available or will become available before the Design Guide is revised again.

Because the aircraft designer must choose the materials for the

aircraft interior, considering all the necessary criteria that

these materials should meet, it is essential that the designer

have the knowledge upon which to base intelligent selections

and trade-offs. Therefore, the following sections present in "

103

.. ~ ~ ~... ....-.... , .• . . .
some detail the various aspects of material flammability hazards, current testing methods, and flammability properties of

some currently used and newly developed materials. Guidelines

for making trade-offs between conflicting criteria also are

presented. This background information should assist the designer in evaluating and selecting interior materials that will

provide maximum fire protection while still meeting necessary design requirements.

5.2 FIRE BEHAVIOR OF MATERIALS

Interior materials can contribute to the overall fire hazard

not only by their flammability, but also by their release of

smoke and toxic gases during combustion. Although the three

factors of flammability, smoke, and toxic gases are discussed

separately in the following sections, all three must be considered together when evaluating any material for its fire

safety.

5.2.1 Flammability

The principal factors to be considered in evaluating the flamnmability of a mnaterial are:

o Ease of ignition.

9 Flame spread rate.

e Heat release rate.

• Flash fire potential.

S5.2.1.1 Ease of Ignition: Ease of ignition can be defined as

the ease with which a material can be ignited under given conditions of temperature, pressure, and oxygen concentration.

Almost any material can be made to ignite with enough heat,

oxygen, and time. Ease of ignition can, therefore, be measured

by the amount of heat required under fixed conditions of oxygen and time, by the amount of oxygen required under fixed conditions of heat and time, or by the amount of time required under

fixed conditions of heat and oxygen.

Ease of ignition can be inferred from minimum radiation intensities required to ignite the material, from the auto-ignition

temperature of the material, or from the minimum amount of oxygen that permits steady burning of the material. These parameters are highly dependent on the conditions under which they

are determined. Test parameters such as sample configuration

and size, ventilation, type of ignition source, superimposed heat input (heat flux), and heat losses can profoundly affect

104

A l I I I
S. . .i ........ . . . . . .. |

the test results. Thus, the relative flammability ranking of

materials may vary with the combustion test used, since a material may perform well in one test and poorly in another.

Generally, the ignition temperature of a material is lower when

the material and the ambient atmosphere are uniformly heated,

as compared to situations in which only the material is heated.

This is illustrated in Table 7, which lists the minimum autoignition temperatures (AIT) obtained in a closed vessel and

the hot plate ignition temperatures in which only the samples

were heated (Reference 33).

TABLE 7. MINIMUM AUTOIGNITION TEMPERATURES (AIT) AND

HOT PLATE IGNITION TEMPERATURES OF SHEETTYPE COMBUSTIBLES IN AIR (FROM REFERENCE 33)

Ignition temperature, *F

Material AIT Hot plate

Cotton sheeting 725 870

Conductive rubber sheeting 735 895

Paper drapes 750 880

Plexiglas sheeting 840 1105

Nomex fabric 960 >1110

Blanket wool 1005 >1110

Cellulose acetate sheeting 1020 >1110

Polyvinyl chloride sheeting 1040 >1110

The time required to ignite a material with a pilot flame is

* dependent on the intensity of any superimposed radiant heat

flux. For instance, under identical test conditions, the time

from flame exposure to burning for particle board varies from

approximately 1.7 min at a radiant heat flux of 0.9 Btu/sec,ft 2

33. Kuchta, J. M., FIRE AND EXPLOSION MANUAL FOR AIRCRAFT ACCIDENT INVESTIGATORS, Pittsburg Mining and Safety Research

Center, Bureau of Mines, AFAPL Technical Report 73-74, Air

Force Aero Propulsion Laboratory, Wright-Patterson Air

Force Base, Ohio, August 1973, AD 771191.

105

*1- '. . , ,. .' l I I I I I I
to 0.5 min at a heat flux of 2.5 Btu/sec,fta (Reference 34).

The minimum oxygen concentration required for combustion also

is dependent on the heat flux seen by the test sample, as shown

in Figure 49 (Reference 35).

The relative flammability hazards of different materials can

be determined for any specific set of test conditions. For

instance, the radiation intensity required for ignition during

tests using a heat flux of 48.7 Btu/sec,ft2 was about 50

Btu/ft 2 for cotton sheeting and between 90 and 120 for wood

and paper sheeting (Reference 33). In comparison, neoprene,

nylon, and polyvinyl chloride sheeting appeared to be nonignitable in air with the same radiatipn source.

Whenever comparisons are made between materials, however, one

must remember that those relative rankings are valid only for

the set of conditions imposed by the test, and may or may not

be valid for other test conditions. The data presented in Figure 49 clearly show the changes in relative rankings that can

occur under varying test conditions.

5.2.1.2 Flame Spread Rate: Surface flame spread can be defined as the rate a flame front travels across a material undea

given conditions of burning. This characteristic provides 

a

measure of fire hazard in that surface flame spread can transmit fire to more flammable materials in the vicinity, thus enlarging the overall fire, although the transmitting material

itself may contribute little fuel to the fire.

Flame spread rates are markedly influenced by such factors as

the presence of a superimposed radiant heat flux, oxygen concentration of the atmosphere, density of the material, and orientation of the material. Generally, flame spread rates increase with increasing radiant heat exposure, as illustrated

in Figure 50 (Reference 35). The magnitude of the change in

flame spread rates can be startling and, at times, misleading

if more than one test condition is not considered. For instance, Smith found that a rigid polyurethane foam that was

self-extinguishing up to a heat flux of 0.5 Btu/sec,ft 2 changed

to a combustible material with a high flame travel rate at 

a

heat flux between 0.5 and 1.0 Btu/sec,ft2 (Reference 34).

34. Smith, E. E., PRODUCT FIRE HAZARD EVALUATION, Journal of

Fire & Flammability/Consumer Product Flammabilfty, Vol. 2,

March 1975, pp. 58-69.

35. Brauman, S. K.,; EFFECT OF SAMPLE TEMPERATURE ON COMBUSTIO14 j• PERFORMANCE OF POLYMERS, Journal of Fire & Flammability,

Vol. 8, April 1977, pp. 210-224.

106 il " 

)
36 0 Polycarbonate

APolyethylene terephthalate

4- Impact polystyrene with

32 alumina trihydrate additive

*-Impact polystyrene with

antimony oxide and

0') chlorinated wax additives

T 28 A-mpact polystyrene

0

4J 0

.- o 24

0 0 Z o

:10

i U2 20

1 2 ,, , , , , !.

Superimposed radiant flux,

2 cal/cm sec

Figure 49. Influence of superimposed heating on

minimum oxygen concentration for

various polymer systems burning in

the driven-rod configuration. (From

Reference 35)

The orientation of the test sample also can markedly influence

flame spread rates. Upward burning of a vertical sample will

* ;generate a higher flame spread rate than will the burning of a

horizontal sample under the same conditions. Also, it has been

observed that the flame spread rate of cotton sheeting is about

40 times greater with upward burning than with downward burning

for specimens in a vertical position (Reference 33).

107

I,.

- .J I
10 0.

Solid symbols represent

full length burn

8.0

6.0

• ~/"

rc

_ Polycarbonate

2.0 2. - OPolyphenylene oxide

Polyethylene

terephthalato

0 Impact. polystyrene

0 0

0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

Radiant flux, cal/cm sec

Figure 50. Influence of superimposed heating

on flame spread rates of various

polymers. (From Reference 35)

5.2.1.3 Heat Release Rate: Heat release can be defined as

the heat produced by the burning of a given weight or volume

of material. This characteristic provides a measure of fire

hazard; i.e., a material that burns with the evolution of little heat per unit quantity burned will contribute less to a

fire than a material that generates large amounts of heat.

Of more importance in relation to the spread of a fire, and

thus, the available esuape time, is the rate of heat release.

Heat release rates can give a comparative measure of the contributions of various materials to a developing fire. However,

108

*,A

, i . . I
I

heat release rate values by themselves cannot adequately describe the contribution of a material in a real fire. In order

to accurately assess the fire hazard of a material, the heat

release rate must be determined as a function of time.

As with the other parameters used to assess a material's flammability, heat release rates are a function of exposure. In

order to predict performance in a fire, heat release data must

be obtained over a range of heat flux levels. Most cellulosic

materials exhibit a uniform change in ignitability, flame

travel rate, and maximum rate of heat release with change in

exposure (Reference 34). However, different materials do not

necessarily respond to changes in exposure in the same manner.

For instance, certain "self-extinguishing" fire-retarded polymers that do not support combustion at low exposure levels will

change to highly combustible materials when exposed to a higher heat flux. This type of behavior can result in two materials,

examined and rated at one set of conditions, having their

ratings reversed at another set of conditions. Wool and nylon

carpet, as well as some of the polymers, go through such rating reversals. At low heat flux levels, nylon is less combustible,

while wool is less combustible at higher heat fluxes (Refer- ence 34).

5.2.1.4 Flash Fire Potential: A flash fire is a flame front

that propagates through a fuel-air mixture as a result of the

energy released from the combustion of the fuel vapor. These

fires occur when combustible vapors evolve from burning materials and accumulate elsewhere as substantial volumes of flammable fuel-air mixtures, which then come in contact with an

ignition source.

Screening tests for the flash-fire propensity of materials have

been proposed based on the concentration of the flammable gases

evolved when the materials are pyrolyzed (Reference 36). The

gases analyzed during the tests were the hydrocarbons methane,

ethylene, ethane, and carbon monoxide. Test results showed

that those materials with the highest propensity for flash

fires, such as polyethylene and polyurethane, had significantly higher hydrocarbon concentrations in their pyrolysis products than did wood, which appeared to have the least propensity for

flash fires. Also, materials that melted, such as polyethylene

and polyurethane foam, had larger concentrations of the more

Sflammable hydrocarbons (ethylene and ethane) than materials

that intumesced, such as bisphenol A polycarbonate, or charred, such as wood.

36. Hialdo, C. J., and Cumming, H. J., HYDROCARBON CONCENTRATIONS IN FIRE TOXICITY TESTS AS AN INDICATION OF FLASH

FIRE PROPENSITY, Journal of Fire & Flammability, Vol. 8,

April 1977, pp. 235-240.

109
5.2.2 Smoke

Combustion of organic materials yields gaseous products in

which small solid particles of carbon and ash, as well as liquid droplets, are frequently dispersed. This mixture of gases,

solids, and liquids can be defined as smoke. In general practice, however, smoke is often defined as the combination of

solid and liquid particles that lead to vision obscuration,

while the gaseous products are treated separately.

The primary hazard of smoke (excluding toxic gases) is the reduction of visibility. The degree of light or sight obscuration due to smoke is generally expressed in terms of optical

density, defined as D - log 100/T (where T - percent light

transmission).

The amount of smoke generated by a burning material depends on

the surface area involved, and the degree of obscuration depends on the available volume and light path length for any given amount of smoke. A quantitative measure, the specific

optical density, has been defined to allow comparisons of smoke

generation between different materials (Reference 37). The

specific optical density is defined as;

D V 100

s a log T (2)

where Ds U specific optical density

•i V - chamber volume all in

A = area of sample exposed to burning consistent S~units

L - path length of light

T - percent light transmission

Ideally, the change in D with time and the maximum D (sometimes designated DM) would depend only on the thickneo of the

material specimen,# its chemical and physical properties, and

the test exposure conditions. The visibility in any size com- .

partment could then be calculated from the D. obtained during

the laboratory testing.

37. Gross, D., et al., SMOKE AND GASES PRODUCED BY BURNING

AIRCRAFT INTERIOR MATERIALS, Building Science Series 18,

U. S. Department of Commerce, National Bureau of Standards, Washington, D. C., February 1969.

110

.X.
a It is difficult to precisely extrapolate specific optical densities to human visibility in burning aircraft compartments. This is because a number of major assumptions must be made in

the extrapolation: the smoke generated is uniformly distributed and is independent of the amount of excess air available;

for any given smoke, the optical density is linearly related

to concentration; and human and photometric vision through

smoke, expressed in terms of optical density, are similar. However, the specific optical density does offer a valid means of

comparing smoke generated by various materials and can be used

to screen out those materials generating the greatest amount of

smoke.

Smoke levels generated by burning materials are dependent on

both physical and chemical parameters of the material involved

and on the burning conditions. In an extensive series of tests

on aircraft interior materials, Gross found that the maximum

smoke level depended on the thickness and density of the specimen and could be expected to increase with thickness, but not

always in direct proportion (Reference 37). He also found that, although most materials produced more smoke during the flaming

exposure test, some materials produced significantly more smoke

in the absence of open flaming (smoldering).

Although the addition of flame retardants has significantly reduced the flammability of many polymeric materials, these chemical additives often have resulted in increased smoke emissions

during fire exposure. Figures 51 and 52 illustrate the effect

of concentrations of reactive and nonreactive flame retardants

on the light obscuration times in rigid urethane foams (Reference 38). It should be noted that the reactive fire retardant,

which imparts the greatest degree of protection, produces more

rapid light obscuration. The addition of flame retardant to a

flexible urethane foam tested by Gross not only resulted in an

increase in overall smoke levels, but also led to a reversal

ot the relative smoke concentrations from smoldering versus

open flaming. 1p

38. Einhorn, I. N., PHYSIO-CHEMICAL STUDY OF SMOKE EMISSION

BY AIRCRAFT INTERIOR MATERIALS, PART I: PHYSIOLOGICAL

AND TOXICOLOGICAL ASPECTS OF SMOKE DURING FIRE EXPOSURE,

University of Utah; Report No. FAA-RD-73-50-1, Department of Transportation, Federal Aviation Administration,

National Aviation Facilities Experimental Center, Atlantic City, New Jersey, July 1973, AD 763602.

S• 111

IDA

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5.2.3 Toxic Gases

The most common gases generated during the combustion of any

organic material are carbon monoxide and carbon dioxide. Several other toxic gases also may be produced, depending on the

chemical composition of the involved material. The results of

the extensive series of burn tests on aircraft interior materials conducted by Gross showed that carbon monoxide (CO) was

produced by almost all the samples in varying amounts depending

on the type of material (Reference 37). Most materials produced significant amounts of other toxic gases in addition to

CO. Table 8 summarizes those results. The addition of flame

retardants can contribute to the generation of taxic gases,

as noted in Table 8, when comparing urethanes identical in all

respects except for the presence of a flame retardant.

TABLE 8. TOXIC GASES PRODUCED BY BURNING

AIRCRAFT INTERIOR MATERIALS

Material CO HCl HCN Other

Nylon X - - -

Wool X - X -

Polyvinyl chloride (PVC) X X - - J

Modacrylic X X X -

Polyamid (aromatic) X - X NO2

Polyvinyl fluoride (PVF) X - - HF

Urethane X - X -

Urethane (flame retarded) X X X -

1

Acrylonitrile/butadiene/

styrene (ABS) X - X -

Polysulfone X - - s02

Rubber X - - So 2

Propy lene X- -

Polycarbonate X -

S113

i ,L
Of course, the amount of toxic gas generated will depend on

the amount of material burned. However, Gross found that the

amount of a given gas produced and its rate of generation are

strongly temperature dependent. This was confirmed during additional testing by Spurgeon, et al., who also found that vary- ing oxygen concentrations will affect the yield of combustion

gases (Reference 39). No generalizations could be made, however, since the observed effects seemed to depend on the composition of the test material. The same is true when comparing

the yields of gases during flaming or nonflaming conditions.

Most of the materials tested by Gross yielded higher concentrations of gases under flaming conditions. There was little difference for some materials, however, while others generated

more gases during nonflaming conditions.

Although approximate human toxicological data are available

for many of the individual gases given off by burning materials, little is known of the synergistic effects of two or more

gases inhaled at the same time. Since the majority of materials give off more than one gas, and since many interior components are actually combinations of materials, relative toxicities of different components must be determined in a manner

that assesses the total effect of the toxic gases given off.

This can be accomplished by using small animal toxicity tests.

In an attempt to correlate analytical test methods with small

animal toxi.city tests, the FAA tested a number of aircraft materials using both methods (Reference 40). Although most

nitrogen-containing materials indicated a correlation between

HCN concentration and time to animal incapacitation, one material (76 percent wool, 24 percent PVC) showed a much higher than

expected toxicity. This toxicity could not be explained on

the basis of HCN concentrations or a simple synergistic response due to the combination of PVC and wool. One possible

k Iexplanation for the observed toxicity is the zirconium fluoride flame-retardant treatment that the material had received.

Whatever the cause, the unexpected toxicity illustrates the

value of animal tests in assigning relative toxicities to interior materials.

39. Spurgeon, J. C., Speitel, L. C., and Feher, R. E., OXIDATIVE PYROLYSIS OF AIRCRAFT INTERIOR MATERIALS, Journal of

?ire & Flammability, Vol. 8, July 1977, pp. 349-363.

40. Spurgeon, J. C., A .'RELIMINARY COMPARISON OF LABORATORY

METHODS FOR ASSIGNING A RELATIVE TOXICITY RANKING TO AIRCRAFT INTERIOR MATERIALS, Report No. FAA-RD-75-37, Department of Transportation, Federal Aviation Administration,

National Aviation Facilities Experimental Center, Atlantic

City, New Jersey, October 1975, AD A018148.

• • 114

1.'4 4

L '.- . --r -'••.•, • l xh '• £ .' '. '",
"H 5.3 MATERIAL TESTING

The number of material tests has increased in direct proportion to the increasing importance placed on fire safety over

the last few years. Unfortunately, the proliferation of tests

has not generated any degrce of consensps in selecting the

"best" test(s) for material screening and selection. There is

a great deal of controversy among those working in this field

as to the validity of the various tests in predicting a material's performance in a real fire. Thus, there are no generally

accepted test methods or crite~iv, for material performance at

the present time. A brief review of the different types of

tests is presented in the following sections so that the designer can understand the various performance ratings assigned

to materials by the use of different tests. Those tests recommended for materials in U. S. Army aircraft are discussed in

Section 5.5. 4 5.3.1 Laboratory Testing

5.3.1.1 Flammability Tests: There are several different types

of tests for material flammability, with each type testing a

specific aspect of flammability, such as ease of ignition,

flame spread, heat release, and fire endurance. In addition,

there are several different test methods for each type of test.

A review of these numerous tests has been compiled by Hilado

(Reference 41).

The simplest tests for ease of ignition provide fixed conditions of heat, oxygen, and time, and the sample either ignites

or does not ignite under those conditions. A somewhat more

sophisticated test is the ASTM D 1929 (Setchkin) ignition test

in which a specimen is exposed to heated air at successively

higher temperatures until ignition occurs. The lowest temperature of air that evolves combustible gases in a sufficient

amount to be ignited by a small pilot flame is defined as the

flash-ignition temperature of the material. The self-ignition

temperature is the lowest air temperature at which the material

ignites by itself, in the absence of any external ignition

source.

S.... A different type of ignition test, one being used increasingly

for aircraft interior materials, is the ASTM D 2863 oxygen index test. In this test, a vertical specimen is ignited at its

upper end by a flame that is then withdrawn; then the atmosphere (mixture of oxygen and nitrogen) that just permits steady

burning is determined. The limiting oxygen index (LOI) is the

41. Hilado, C. J., FLAMMABILITY TESTS, 1975: A REVIEW, Fire

Technology, Vol. 11, No. 4, November 1975, pp. 282- =9.

115

................................
minimum concentration of oxygen in the oxygen-nitrogen mixture

that will just permit the sample to burn. The higher the LOI,

the less flammable the material is.

In addition to the above two widely used tests, Hilado lists

eight other tests for ease of ignition. He points out that

many tests that measure flame spread are actually tests for

ease of ignition because failure to ignite or to sustain ignition is the most desirable response. Hilado lists 10 tests

in this category, including the FAR 25.853 vertical test (Reference 42).

The latter test is currently required by the FAA for compartment interior materials in transport category airplanes. In

this test, the lower edge of a vertically mounted sample is

exposed to a burner flame for either 60 or 12 sec, depending

on the type and application of the material. The flame is then

removed and flame time, burn length, and flaming time of drip- pings are recorded. All materials must be self-extinguishing;

i.e., average flame time after removal of the flame source must

not exceed 15 sec. In addition, the average burn length must

not exceed 6 or 8 in., again depending on the type and applica- tion of materials.

There are even more tests for surface flame spread rates;

Hilado lists 35 tests in this category, including the FAR

25.853 test. Two of the most widely used tests for flame

spread are the ASTM E 84 tunnel test and the ASTM E 162 radiant panel test.

The tunnel test involves igniting the sample at one end and

measuring the rate at which the flame travels across the surface. Flame spread classification is determined on a scale

where asbestos cement board is zero and select-grade red oak

flooring is 100. No additional heat flux is supplied beyond

that generated from the burners and the burning sample itself.

The radiant panel test, however, does provide additional heat

beyond that generated by the sample. The specimen is mounted

in a semivertical position in front of a radiant heat source

maintained at 6700C (12380F). The top of the specimen is then

ignited and the flame spread index determined as in the tunnel test.

Many authorities feel that the radiant panel test is superior

because it more nearly duplicates the exposure the material

would experience in an actual fire. (The importance of the

increased heat exposure in assessing a material's "true" flammability has been discussed in Section 5.2.1.) The Urban Mass

42. Federal Aviation Regulations, AIRWORTHINESS STANDARDS:

TRANSPORT CATEGORY AIRPLANES, Part 25, Section 25.853.

116
Transportation Administration (UMTA) has recommended the radiant panel test in their guidelines for flammability 'specifica- tions applicable to combustible materials used in transit systems (Reference 43). The National Materials Advisory Board of

the National Research Council also has recommended the use

of ASTM E 162 for aircraft interior materials testing (Reference 44).

There are several tests for heat release, but currently they

are not as widely used as the more established tests for ease

of ignition and flame spread. The Factory Mutual calorimeter

test and the Ohio State University heat release rate tests are,

the most commonly used at the present time.

Fire endurance tests measure the resistance offered by a material to the passage of fire normal to the exposed surface. The

fire resistance can be measured by the burn-through time or by

the relative difference in temperature between the flame side

and the back face of the specimen. There are several tests

for fire endurance. The NASA Ames T-3 test (Reference 45)

seems to be the most widely used endurance test for aircraft

materials and is well suited for testing components containing

several different materials, such as interior fuselage wall

panels.

5.3.1.2 Smoke Evolution Tests: One of the earliest procedures for measuring smoke density was the ASTM D 2843 test.

This test measures the light obscuration over a 1-ft optical

path inside an enclosed chamber containing the burning sample

(Reference 38). Smoke evolution also can be measured during

flammability testing in the ASTM E 84 tunnel test, the ASTM E

162 radiant panel test, and the heat release test developed at

Ohio State University. The most widely used test for aircraft

materials, however, is the National Bureau of Standards (NBS)

smoke density test first developed by Gross, et al. (Reference 37).

43. Transportation Systems Center, PROPOSED GUIDELINES FOR

FLAMMABILITY AND SMOKE EMISSIONS SPECIFICATIONS, (Unofficial) U. S. Department of Transportation, Cambridge,

Massachusetts.

44. FIRE SAFETY ASPECTS OF POLYMERIC MATERIALS, VOLUME 6 -

AIRCRAFT: CIVIL AND MILITARY, National Materials Advisory

Board, National Academy of Sciences, Washington, D. C.,

1977, p. 11.

45. Fish, R. H., AMES T-3 FIRE TEST FACILITY - AIRCEAFT CRASH

FIRE SIMULATION, Journal of Fire & Flammability, Vol. 7,

October 1976, pp. 470-481.

117

!,,
The NBS test, conducted in a completely closed cabinet, exposes

a vertical sample to 2.5 W/cma (2.2 Btu/sec,ft 2 ) thermal radiation from an electric heater. Light absorption is measured

by a photometer over a vertical light path 3 ft long. Tests

are performed under both flaming and nonflaming (smoldering) conditions. (A small pilot flame applied to the bottom of the

specimen induces open flaming.) Smoke measurements are expressed in terms of specific optical density, D (refer to Section 5.2.2). Both the FAA and UMTA have proposid the NBS smoke

density test for interior materials (References 43 and 46).

The proposed FAA requirement specifies that the D of interior

materials must not exceed 100 in 4 min for textills, air ducting, thermal insulation, and insulating covering. The D of

all other interior materials shall not exceed 100 withinS 90

seconds, nor 200 within 4 min.

5.3.1.3 Toxic Gas Tests: Concentrations of potentially toxic

gases can be determined by chemical analysis of the combustion products. However, the results of this analysis depends on

the accuracy of the analytical method employed, the effectiveness of the sampling technique used, and the number of different gases analyzed. Even though the previous factors might be

optintized, the problem of relating the chemical results to

physiological hazards still remains. Synergistic toxicological

effects of various gas combinations are not amenable to analysis. The possibility also exists that some toxic components will not be anticipated and, therefore, will not be considered

in the analysis. This latter situation arose during FAA tests

comparing chemical analysis versus animal toxicity tests in assigning relative toxicity hazards to aircraft materials (Reference 40).

Both the FAA and the University of San Francisco, under NASA

sponsorship, have done extensive animal toxicity testing of

•! 1

46. Federal Aviation Administration, Proposed Regulations, TRANSPORT CATEGORY AIRPLANES: SMOKE EMISSION FROM

COMPARTMENT INTERIOR MATERIALS, Federal Register, 40

F.R.6506, 12 February 1975.

118

0 f
aircraft materials (Reference 47 and 48). Both laboratories

expose rodents (the FAA uses rats, USF uses mice) to the pyrol- ysis products of materials that are thermally degraded in a

tube furnace. The exposure conditions vary, however, since

the FAA tests maintain a nearly normal concentration of oxygen

in the exposure chamber, while the USF tests do not. The FAA

conducts its tests with sufficient ventilation in the pyrolysis furnace to assure near normal oxygen concentrations, while USF

runs its tests with or without air flow through the furnace.

Both laboratories report times to incapacitation and times to

death.

Although the above tests cannot duplicate the gas concentrations found in a real fire, the tests do reflect the relative

toxicities of different materials to rodents under the specific

test conditions. Thus, the animal toxicity tests are useful

in screening out the most hazardous materials. The FAA also

found that dose-response relationships for the systemic toxins

(CO, HCN) are very similar for rodents and humans (Reference 47).

5.3.2 Large-Scale Testing

Because the validity of laboratory tests as a means of predicting material behavior in a real fire has been of increasing

concern, large-scale testing is being used more and more as a

final test for system performance in a fire. Part of the problem with the predictability of the laboratory tests lies in

the fact that a system of materials, not just one material, is

involved in actual fires. The types of materials, amounts of

each material, and the locations of the different materials

all affect the development of a fire in an aircraft fuselage. The behavior of one material will affect that of another, possibly altering its behavior markedly from that demonstrated in

laboratory tests.

Large-scale tests of bus and rail car interior assembly mockups

illustrate the total effect of all the materials comprising a

47. Crane, C. R., Ph.D., et al., INHALATION TOXICOLOGY: I.

DESIGN OF A SMALL-ANIMAL TEST SYSTEM: II. DETERMINATION

OF THE RELATIVE TOXIC HAZARDS OF 75 AIRCRAFT CABIN MATERIALS, FAA Civil Aeromedical Institute; Report No. FAA-AM77-9, Department of Transportation, Federal Aviation Administration, Office of Aviation Medicine, Washington,

D. C., March 1977, AD A043646.

"48. Hilado, C. J., FIRE RESPONSE CHARACTERISTICS OF NONMETALLIC MATERIALS: A REVIEW OF RECENT PAPERS AND REPORTS,

Journal of Fire & Flammability, Vol. 7, October 1976,

pp. 539-558.

119

I4
system (Reference 49). The assemblies consisted of one or two

seat assemblies, wall paneling, and glazing, as would be found

in the actual vehicle. Effects of various seat cushions, seat

backs, an4 glazing materials were studied during fire tests

started by igniting newspapers on the seat. The tests showed

that with urethane seat cushions the system failed at approx- imately 6 min (flashover occurred) irrespective of the glazing

or wall-covering material. Replacing the urethane with less

flammable neoprene cushions resulted in increasing the importance of other materials in the system performance. For example, acrylic glazing panels led to system failure (total involvement near flashover conditions) at 7 min, even with the

presence of the neoprene seats. However, the fire was confined

to the seat of origin in tests using neoprene seat cushions

and polycarbonate glazing.

McDonnell Douglas Corporation has developed a Cabin Fire Simulator (CFS) for use in testing commercial aircraft interiors

(Reference 50). The CFS is a double-walled steel cylinder

12 ft in diameter and 40 ft long, equipped with a ventilation

system, exhaust scrubber, and nitrogen-extinguishing system. To date, over 200 individual fire tests have been conducted in

the CFS. These tests have been valuable not only in assessing material interactions but also in evaluating design changes.

For instance, in a full cabin lavatory fire test, the lavatory module failed to contain the fire and the fire erupted into

the cabin area. Analysis showed that the failure to contain

the fire was primarily due to a utility panel falling from the

ceiling, resulting in the escape of combustible gases into the

cabin. Covering the panel considerably improved system performance.

5.4 SELECTED MATERIAL PROPERTIES

Until recently, most efforts to reduce the flammability of ma- : terials were centered on flame-retardant treatments of polymers with halogens or phosphorous. This led to materials such as

Fluorel, Refset, and Durette, as well as flame-retarded polyurethanes and cottons, wools, and other fabrics. Many recent

!; efforts, however, have been centered on developing thermally stable, char-forming polymers such as polyimide, polyphosphazene, and polybenzimidazole (PBI). Interest in the thermally

49. Nelson, G. L., et al., MATERIAL PERFORMANCE IN TRANSPORTATION VEHICLE INTERIORS, Journal of Fire & Flammability, Vol. 8, July 1977, pp. 262-77F.

50. Dusken, F. E., FIRE TESTING OF AIRCRAFT CABINS, Journal

of Fire & Flammability, Vol. 8, April 1977, pp 1W3-2-01

120

... ....... ,

S"a
stable polymers has increased with the finding that most flameretarded materials produce significantly greater amouDas of

smoke and toxic gases than do the thermally stable polymers.

Although the listing of flammability properties for the numerous materials already being used or being deve~oped for use in

transportation vehicle interiors is beyond the scope of this

volume, a brief overview of some current and newly developed materials for aircraft interiors follows.

5.4.1 Seat Cushion Foams

Polyurethane foam is the most common seat cushioning material

currently used for aircraft seats. Studies conducted in the

late 1960's showed that fire-retarded polyurethane foam was considerably less flammable than nontreated foam (Reference 51).

Since that time, considerable efforts have been expended in

trying to improve the flame resistance of polyurethane. Einhorn concluded, after his studies, that major improvements

could be accomplished in the flammability characteristics of

rigid polyl.iethane foams by modifying the chemical structure

and formulation (Reference 52). However, the flexible foam

system did not possess the necessary chemical structure to permit the formulation of truly flame-resistant systems.

Einhorn's conclusions seem to be validated by the results of

NASA tests that exposed fire-retardant-treated, rluorel-coated,

and untreated, uncoated polyurethane foarmi seati to a large

flaming ignition source located 12 in. below the seat cushion

(Reference 53). These test results indicated that the improved

51. Marcy, J. F., and Johnson, R., FLAMING AND SELF-EXTINGUISHING CHARACTERISTICS OF AIRCRAFT CABIN INTERIOR MATERIALS, Report No. NA-68-30 CDS-68-13], Department of

Transportation, Federal Aviation Administration, National

Aviation Facilities Experimental Center, Atlantic City,

New Jersey, July 1968, AD 673084.

52. Einhorn, I. M., Kanakia, M. D., and Seader, J. D., PHYSIOCHEMICAL STUDY OF SMOKE EMISSION BY AIRCRAFT INTERIOR MATERIALS, PART I1: RIGID- AND FLEXIBLE-URETHANE FOAMS,

University of Utah; Report No. FAA-RD-73-50,II, Department

of Transportation, Federal Aviation Administration, National Aviation Facilities Experimental Center, Atlantic

City, New Jersey, July 1973, AD 763935.

53. Fewell, L. L., et al., eds., CONFERENCE ON THE DEVELOPMENT

OF FIRE-RESISTANT AIRCRAFT PASSLNGER SEATS, NASA Technical

Memorandum X-73144, Ames Research Center, National Aeronautics and Space Administration, Washington, D. C., August 1976, pp. 115-125.

l121

S t÷ L • " '" . . '
state-of-the-art polyurethane foams without the added fire retardant and coating treatments were not significantly better

than untreated, older, less fire-resistant foams. However, by

treating and coating the state-of-the-art foams, production of toxic gases was delayed, and destruction of the foam was limited. It should be noted that relatively high levels of hydrogen cyanide were detected in each test, indicating that poly- j urethane foam may be the major contributor to similar high

levels found in large-scale tests. Figures 53 and 54 show the

temperatures of the top portions of the seat backs and the hy- drogen cyanide concentrations during the tests.

Neoprene foam, used extensively in mattresses, has been advocated for seat cushions and is currently being used for that

purpose in some mass transit vehicles. Its flammability characteristics are superior even to flame-retarded polyurethane

foam. One of the earliest large-scale tests comparing neoprene foam with other seat cushioning materials was conducted by the

FAA using a simulated airplane cabin. The test data comparing

urethane and neoprene foam seat cushions are summarized in

Table 9 (Reference 54). These tests were some of the earliest

studies to document the now well-known flashover phenomenon encountered with unretarded urethane foam. Although the flame- " retarded urethane foam was effective in reducing fire temperatures, neoprene was more effective. The smoke levels in these

tests showed that neoprene provided significantly longer times

to 50 percent smoke obscuration, a result repeated in large- scale bus and rail car tests (Reference 49).

The smoke results in the large-scale tests of neoprene foam contrast markedly with NBS smoke chamber data that indicate neoprene releases more smoke than urethane. This anomaly illustrates many of the problems in trying to extrapolate laboratory

data to real fire situations. Under actual fire conditions,

neoprene decomposes at a slower rate than it does in the NBS

laboratory test, thus producing smoke at a lower rate. University of San Francisco tests also have shown that the neoprene

foam produces less toxic smoke than-does the polyurethane foam

(Reference 55). Efforts now are under way to develop improved

neoprene foam formulations that have superior smoke evolution

properties. Table 10 summarizes comparative data on the standard and improved neoprene foams.

54. Morford, R. H., THE FLAMMABILITY OF NEOPRENE CUSHIONING

FOAM, Journal of Fire & Flammability, Vol. 8, July 1977,

pp. 279-299.

55. Hilado, C. J., THE CORRELATION OF LABORATORY TEST RESULTS

WITH BEHAVIOR IN REAL FIRES, Journal of Fire & Flammabilit Vol. 8, April 1977, pp. 202-209.

122

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TABLE 9. RESULTS OF LARGE-SCALE AIR TRANSPORT CABIN SFIRE TESTS USING URETHANE OR NEOPRENE SEAT

CUSHIONS (FROM REFERENCE 54)

Test number

Property 4 11 22A 22B

Seat cushion Urethane.F.R. urethane Neoprene Neoprene

material foam foam foam foam

Maximum ceiling

temperature, OF 1420 560 230 100

Time to 50 percent

smoke obscuration,

min 1.9 1.7 3.4 12.4

Minimum oxygen

concentration,

percent 2.5 20.0 18.3 20.5

Maximum CO concen-"

tration, percent 1.5+ 0.6 0.21 0.05

Time to flashover, 2.6 0 0

min (4 min)

TABLE 10. TYPICAL PROPERTIES OF STANDARD AND IMPROVED

NEOPRENE FOAM (FROM REFERENCE 54)

Standard 7 oroved

Property foam Zoam

Density, lb/ft 3 4 7

Tensile strength, lb/in. 2 9 7

Elongation, percent 120 100

Compression set

(50 percent deflection, 22 hr @ 212*F) 7.5 7.5

ASTM E-162-75 flame spread rating 6 4

NBS smoke chamber results

D ,90 sec 280 93

D 4 min 380 195

D , maximum 380 250

T~me to maximum D min 4 7

( 125
* Advanced state-of-the-art materials being developed for seat

cushions include polyphosphazene and polyimide foams (References 56 and 57). Both of these foams have superior flammability properties compared to typical fire-retarded urethane foams,

as can be seen in Table 11. Future optimization and production

of these types of foams may make them viable alternatives for

improved aircraft seat cushions.

TABLE 11. COMPARISON OF POLYPHOSPHAZENE, POLYIMIDE, AND

FIRE-RETARDED POLYURETHANE FOAM PROPERTIES

Typical Typical Typical polyphosphazene polyimide F.R. urethane

Property foam foam foam

Density, lb/ft 3 4.0 - 9.0 1.3 - 1.4 4.5 - 8.5

Tensile strength,

psi 20 - 80 10 - 13 40

Elongation, percent 80 - 125 20 - 23 100

Flame spread index 14 30

Limiting oxygen

index 43 - 45 44 -54 20

Maximum smoke

density, Ds

Flaming 40 - 150 0 - 0.5 250

Nonflaming 0 - 1 -

56. Parker, J. A., Ph.D., POLYPHOSPHAZENE SEAT CUSHION APPLICATIONS, American Research Center; in Conference on the

Development of Fire-Resistant Aircraft Passenger Seats,

Fewell, L. L., et al., eds., NASA Technical Memorandum

X-73144, National Aeronautics and Space Administration,

Washington, D. C., August 1976.

57. Gagliani, J., FIRE RESISTANT RESILIENT FOAMS - FINAL REPORT, Solar Division, International Harvester; Report No.

N76-18278, National Aeronautics and Space Administration,

Lyndon B. Johnson Space Center, Houston, Texas, February

1976.

1 126

. ... . . . . . . ... . . ... . . ...... 

S. .. . , n i n ! I I I I
Another recent approach to improving the fire resistance of

aircraft seats is the use of an interliner or thermal barrier

between the upholstery fabric and the seat cushion foam (Refererence 58). However, experimental evaluation of specific combinations of materials on systems is essential if this approach

is used.

5.4.2 Upholstery and Other Fabrics

Current seat upholstery and other fabrics can be divided into

two classes: uncoated and coated. Typical uncoated fabrics

include wool, cotton, nylon, rayon, polyester, modacrylic, or

combinations of these fibers. Marcy found that the majority

of these fabrics, even when treated with .Lame retardants, had

unacceptable burn rates during both the FAA vertical flammability test and the ASTM radiant panel test (Reference 51). Modacrylic fabrics were the only self-extinguishing fabrics. However, animal toxicity tests revealed that modacrylic fabrics

were the most toxic of 75 different interior materials that

were tested (Reference 47).

The coated fabrics, on the other hand, were all self-extinguishing during the vertical burn tests. These materials consisted

in large part of vinyl- and acrylic-coated glass fabrics. Toxicity tests of cotton, nylon, and polyester fabrics coated with

polyvinyl chloride (PVC) showed that these coated fabrics were

much less toxic than their uncoated counterparts.

Advanced state-of-the-art fabrics include Nomex, Kynol, polybenzimidazole (PBI), and polyimide (e.g., Kapton) fabrics.

These fabrics exhibit superior. flammability characteristics,

as shown in Table 12 (from Reference.59). Kynol and PBI fabrics emit very little smoke and essentially no toxic gases during burning. Nomex is readily available, while Kynol and PBI

58. Tesoro, G. C., Ph.D., FABRICS FOR FIRE RESISTANT PASSENGER

SEATS IN AIRCRAFT, Massachusetts Institute of Technology;

in Conference on Fire Resistant MaLerials (Firemen),

Kourtides, D. A., ed., NASA Technical Memorandum 78523, National Aeronautics and Space Administration, Ames Research

Center, Moffett Field, California, October 1978.

59. Tesoro, G. C., Ph.D., STATE OF THE ART MATERIALS USED IN

AIRCRAFT PASSENGER SEATS, Massachusetts Institute of Technology; in Conference on the Development of Fire-Resistant

Aircraft Passenger Seats, Fewell, L. L., et al., eds., NASA

Technical Memorandum X-73144, National Aeronautics and

Space Administration, Washington, D. C., August 1976.

127
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Scurrently are subject to limited availability. However, avail- 4 ability of the latter two fabrics should increase as demand

increases.

5.4.3 Structural Components

A.ll major transport aircraft manufacturers and NASA are en- gaged in efforts to increase the fire safety of interior struc- tural components, such as sidewall, floor, and ceiling panels.

These efforts encompass the selection of single candidate materials and the fabrication of multimaterial assemblies.

* Candidates for improved thermoplastic materials are listed in

Tables 13 and 14, along with their physical and chemical properties (Reference 60). All of these materials exhibit greater fire resistance and lower smoke and toxicity than the majority

of aircraft interior materials currently in use. Other promising materials include the char-forming polyisocyanurate and

PBI foams (Reference 61). Fire-hardening of honeycomb panels

has been accomplished by filling the honeycomb core with PBI

or isocyanurate foam, as well as with phenolic-impregnated fiberglass batting (References 50 and 62). Further improvements

can be made by replacing flammable adhesives (e.g., acrylate

adhesive) with more fire-resistant compounds such as fireretarded epoxy adhesives or polyamide adhesives (Reference 61).

5.4.4 Aircraft Insulation

Suitable insulation installed in the interior walls of crew

and passenger compartments can help protect occupants from heat

60. Silverman, B., AIRCRAFT INTERIOR THERMOPLASTIC MATERIALS,

Lockheed Missiles and Space Company; in Conference on the

Development of Fire-Resistant Aircraft Passenger Seats,

eds. Fewell, L. L., et al., NASA Technical Memorandum

X-73144, National Aeronautics and Space Administration,

Washington, D. C., August 1976.

61 Parker, J. A., Ph.D., et al., FIRE DYNAMICS OF MODERN.

AIRCRAFT FROM A MATERIALS POINT OF VIEW, in Aircraft Fire

Safety, AGARD Conference Proceedings No. 166, North Atlantic Treaty Organization, Advisory Group for Aerospace Research and Development, Neuilly-sur-Seine, France, October

1975, pp. 10.1-10.11, AD A018180.

62. Ander'qon, R. A., et al., EVALUATION OF MATERIALS AND CONCEPTS FOR AIRCRAFT FIRE PROTECTION, Boeing Commercial Airplane Company; Report No. NASA CR-137838, National Aeronautics and Space Administration, Ames Research Center,

Moffett Field, California, April 1976.

S( 129

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generated from an exterior postcrash fire as long as the fuselage remains intact. However, not all types of insulation are

suitable, and some insulations, such as glass wool or fiberglass, may worsen, rather than lessen, the postcrash fire hazard.

Recent developments in foam insulation may offer promising results for the aircraft industry. Tests conducted under the

sponsorship of the NASA/Ames Research Center provided evidence

of the effectiveness of a modified polyisocyanurate semirigid

foam. A C-47 aircraft fuselage was separated into two zones:

one a reference zone with no modification, and the other, a

zone provided with the foam insulation, as shown in Figure 55.

Fuel-fed ground fires were lighted next to the aircraft fuselage and allowed to burn for approximately 10 min. Temperatures recorded inside the aircraft fuselage are shown in Figure 56. Little smoke or gas was evolved. Occupant survival

time inside -e reference section would have been no more than

1-1/2 min, while an occupant in the insulated section could

have survived for approximately 9 to 10 min. (See Section

3.3.1 for a detailed discussion of human tolerance to heat.)

The application of similar technology to rotary-wing aircraft

has not been as successful. Fire tests of a CH-47 and a UH-lD

helicopter showed that only limited protection could be attained for occupants for the first few minutes following the

onset of a postcrash fire (Reference 63). Two main factors

accounted for the limited protection attained. These were (1)

the unreliability of the protective wall materials because of

problems in suitably applying the materials to the helicopter

wall structures, and (2) the poor fire resistance of currently

used Plexiglas helicopter windows.

The walls of the two helicopters posed different insulation

problems because of their different structures. Fire penetrations in the CH-47 walls occurred where the isocyanurate foam

could not be applied because of the presence of wiring, air

ducts, and hydraulic oil tubes. The UH-ID walls did not lend

themselves to foaming because of the absence of ribs and

formers. The sodium silicate hydrate panels used to protect

the interior walls partially collapsed because of the absence

of structural support.

63. Atallah, S., and Buccigross, H., INVESTIGATION AND EVALUATION OF NONFLAMMABLE, FIRE-RETARDANT MATERIALS, Arthur

D. Little, Inc.; USAAMRDL Technical Report 72-52, Eustis

Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, November 1972,

AD 906699L.

132 (

U .', N, W
1, il . ,•Selected test area I

I, , __ -- - - --

...... L -Protected zone

Insulated steel Reference zone barriers (3)

C-47 aircraft fuselage I -- ~Fuselage! • shell Stringers /

2.5 in.

of foam

Fuselage" 2-3 in. skin Intumescent ,

Of fcam paint

floor sprayed

14 1 board to 2.5 in.

thick

Single layer (full

boards loorof matting fiberglass depth of bonded frame).

to skin 1/32 in.

Section fiberglass through General epoxy sheet protected zone construction

of foam

Details of foam installation

Figure 55. Semirigid foam insulation test configuration.

133

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-- .: - ... ••,.•w•,•W.•!"',i,:
500

Reference section

400 .,

0

300

4J

ý4

SInsulated section--

i00 -

0 1. 2 3 4 5 6 7 8 9 10

Time, min

Figure 56. Compartment ambient temperature versus time from

fuel ignition during foam insulation tests.

Although the postcrash fire protection was limited, in-flight

simulation tests conducted during the same program indicated

that it should be possible to protect the habitable compartment against a fire occurring in an adjacent compartment. Sodium

silicate hydrate panels lining the fire compartment successfully contained the fire and kept temperatures in the adjacent

cabin far below human tolerance levels.

134

L
Intumescent paints and coatings provide another method of thermal insulation. These fire-retardant materials react to the

heat of a fire by swelling and forming a thick, low-density,

polymeric coating or char layer, thus protecting the coated

surface from the full effects of the fire. An intumescent

paint was sprayed on the interior fuselage wall in the CH-47

foam test described above to provide a void fill between the

frame-foam interface to prevent burn-through at that point.

More research must be done in this area, however, before these

materials can provide postcrash fire protection by themselves.

Furnace test results for eight commercially available intumescent paints and coatings indicated that none of them could provide the desired fire protection (Reference 63). Although

these materials intumesced readily, the fire gases eroded the

char very quickly when the coatings were exposed directly to

the fire. The coatings were not effective on the nonfire side

of the aluminum panel because the char could not support itself

once the panel burned away. Furthermore, most of them produced

noxious fumes, and therefore should not be applied in habitable

compartments. These materials did show promise, however, as

linings for potential fire compartments to protect occupants

against in-flight fires resulting from fuel or hydraulic line leakage. Flight-critical components next to a fire zone can also receive partial protection from these paints.

5.5 MATERIAL SELECTION CRITERIA

5.5.1 General Considerations

As stated in the introduction to this chapter, the selection

of interior materials is governed by several varied, and sometimes conflicting, requirements. All interior materials must

effectively meet their original intended uses. For instance,

seat cushions must provide a certain degree of comfort, must

possess specific crashworthy characteristics of compressive

modulus an• rebound, and must meet minimum durability criteria..

The additional requirements of low flammability and low smoke

and toxic gas emissions can create a problem in finding materials that meet all of the criteria, and trade-offs must sometimes be made. The following guidelines should be considered

in establishing priorities for trade-offs.

The importance of material flammability characteristics increases as the fuel load of any material(s) in the aircraft inJ." creases. Those materials most prevalent in the aircraft, such

as noise insulation and interior structural components, should

possess as low a flammability rating as possible. Low flammability characteristics for restraint systems, on the other

135

A /
hand, are not as critical because of the small amount of material used. Seat cushions and upholstery, depending on the

quantity used, might fall somewhere in between in regard to

the importance of their flammability characteristics. This

should not be construed as deleting flammability requirements

for those materials used in lesser quantities. Every effort

should be made to select the least. flammable material available

for each end use.

Thope items involved directly in crashworthiness, such as seat

cushions and restraint systems, must satisfy all crashworthy

requirements as a first priority. If there is no material

available that can satisfy both the crash and postcrash (flammability) requirements, some reduction in optimum flammability characteristics might have to be tolerated. However, the designer should first consider protecting the more flammable material with a less flammable one. The effectiveness of the

latter approach must be confirmed by flammability tests of the

candidate system to ensure that the system performs as antici- pated.

If materials that cannot fulfill all of the flammability requirements contained in the following section must be selected,

materials that present the least amount of fire hazard should

be chosen. The material should not ignite easily and should

have as low a flame spread rate as possible. Care must be

taken to avoid selecting a material with high flashover potential, such as a nonfire-retarded polyurethane foam.

Smoke and toxic gas emissions also should be held to minimum

possible levels. Those materials known to emit significant

amounts. of toxic gases, such as modacrylics, should not be

used.

5.5.2 Flammability Test Criteria

The large number of laboratory tests available, plus the controversy among experts as to which test or tests are preferred,

has precluded the adoption of one, or even one set of tests

for all materials. There are, in fact, several tests recommended for vehicle interior materials.

At the present time, the FAA flammability requirements specified in FAR 25.853 (Reference 42) are the only specific, mandatory requirements for aircraft interior materials. The FAA

proposed standard for smoke emission has not yet been adopted.

In 1977, the FAA held extensive hearings on the feasibility of

developing a comprehensive set of integrated flammability,

smoke emission, and toxic gas emission standards for interior

Itiaterials in transport category aircraft. Currently, the

matter is under study, and no recommendations have been issued.

136

- .. . ... . . ... . . .. .. .. . I.
Recommended specifications for transit system combustible materials have been proposed by UMTA (Reference 43). These spe-

* cifications are guidelines only and do not have any official

status. However, many manufacturers in the transit industry

are voluntarily following these guidelines. The guidelines establish specifications for both flammability and smoke emission, but none, as yet, for toxic gas emission.

Although many of the other tests can and should be used for

screening materials during the selection process, until some

agreement is reached among experts and regulators as to their

suitability for standards, U. S. Army criteria should he based

on the FAA requirements or the UMTA guidelines. As a minimum,

interior materials in all U. S. Army aircraft must meet the requirements for FAR 25.853. Those aircraft that can carry passengers (i.e., utility and cargo aircraft as opposed to attack

or observation aircraft) should meet the more stringent guide- lines for both flammability and smoke emission issued by UMTA.

5.5.2.1 eAR 25.853 Flammability Requirements: Materials used

in each ctrmpartment occupied by the crew or passengers must

meet the following requirements:

* Ceiling panels, wall panels, partitions, structural

flooring, etc. Must be self-extinguishing when

tested vertically by applying a 15500F flame to the

lower edge of the specimen for 60 sec. Average burn length not to exceed 6 in.; average flame time after

removal of test flame not to exceed 15 sec. Drip- pings may not continue to flame more than an average of 3 sec.

* Floor coverings, textiles (including upholstery),

seat cushion, adigs,inlaos eeŽitrical insulation), etc. Must be self-extinguishing when tested vertically by applying a 15506F flame to

J the lower edge of the specimen for 12 sec. Average burn length not to exceed 8 in., average flame time

after removal of test flame not to exceed 15 sec.

Drippings may not continue to flame more than an

average of 5 sec.

SAcrylic windowe, signs, restraint systems, etc. May

not have an average burn rate greater than 2.5 in./

min when tested horizontally by applying a 1550*F

flame to the specimen edge for 15 sec,

The reader is referred to Reference 44 for the complete text of

tlie regulations and test requirements.

137

Sf o
5.5.2.2 UMTA Flammability and Smoke Emission Guidelines: Combustible materials used in transit systems are required to

possess the following flammability characteristics:

B Seat cushions and insulations (except electrical

irklation. Must pass ASTM E 162-67 Radiant Panel

Test with a flame propagation index (I ) not exceeding 25, with the added provision that fhere shall be

no flaming, running, or dripping.

* Wall and ceiling panels, seat frames, partitions,

etc. Must pass ASTM E 162-67 Radiant Panel Test with

fa-flame propagation index (I ) not exceeding 35, with

the added provision that thire shall be no flaming

dripping.

Upholstery materials. Burn length must not exceed

6 in. when tested by FAR 25.853 vortical test. Average flame time after removal of flame source may not

exceed 10 seu. Flaming dripping not allowed.

e Carpeting (tested with its padding) Must pass NBS

flooring Radiant Panel Test, NBSR-74-495, with a

minimum critical radiant flux of 0.6 W/cma.

a Plastic windows. Must pass ASTM E 162-67 Radiant

Panel Test with a flame propagation index (Is) not

exceeding 100.

* Flooring. Must withstand requirements of ASTM E 119

when underside is exposed to a flame up to 1400OF for

15 min.

* Elastomers. Must pass the requirements of ASTM C542-

"A, with the added requirement of no flaming drip- ping.

When tested in accordance with the National Fire Protection

AAssociation Standard No. 258 (1974) in both flaming and nonflaming modes, combustible materials should meet the following smoke emission requirements:

* Upholstery, air ducting, insulation (except electriCal insulation). Optical density (D ) must not exceed 100 within 4 min after start of test.

138 i' - '
" All other materials, (except foam seat cushioning, electrical insulation, and carpeting). Optical denA", sity (D) must not exceed 100 within 90 sec after

start of test, nor exceed 200 within 4 min after

start of test.

If fire-retardant coatings are used for fabric and trim materials, the effects, if any, of routine maintenance and cleaning procedures must be assessed. If the coatings can be removed

by routine cleaning procedures, the flammability and smoke/

toxic fume tests should be repeated after a representative

* number of cleaning cycles.

The above guidelines are being voluntarily used by several

transit authorities and manufacturers although the guidelines

are not Government standards and have no official status. The

reader is referred to Reference 43 for the complete text of

the regulations and test requirements.

5.5.3 Evaluation of Materials

The flammability properties of polymeric materials currently

used in commercial aircraft have been qualitatively assessed

by the Committee on Fire Safety Aspects of Polymeric Materials

of the National Materials Advisory Board (Reference 44). This

information is presented in Table 15 to assist the designer in

evaluating materials. Table 15 should not be construed as a

list of acceptable materials, since the assessment is based on

flammability properties only. Specific application to U. S. Army aircraft, which necessarily have different functional requirements than do commercial aircraft, must be considered in

final material selection.

Also, a great deal of effort is being expended in the development of newer, less flammable, less toxic interior materials.

The flammability properties of these materials are scattered

throughout the literature. Fortunately for the designer, the

Transportation Systems Center (TSC) of the U. S. Department of

Transportation has established a computerized materials information bank that will be continually updated as new materials

and testing methods are evaluated. The data bank contains flammability, smoke, and toxicity properties obtained by a variety

of standard testing procedures for candidate materials. These

data are supplemented with available physical and mechanical

properties, as well as durability and maintainability data.

The data in the bank can be accessed only from TSC, located in

Cambridge, Massachusetts. The data will be made available to

designers, transit authorities, and other Government agencies.

139

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S-'" 6. DITCHING PROVISIONS

____6.1 INTRODUCTION_ _

Since U. S. Army aircraft are frequently flown over water, un-

* planned water landings are not uncommmon. The crash parameters, survival envelope criteria, and human tolerance limits

presented in Volume 2 of this design guide are equally applicable to water and land impacts. However, the water environment during the postcrash phase presents additional unique

problems that do not occur on land. This chapter addresses

those problems and offers general design concepts and guidelines to increase occupant survival in ditching situations.

6.2 THE DITCHING ENVIRONMENT

An aircraft ditching is a forced landing of an aircraft in the

water. It is not to be confused with an uncontrollable crash

into a water environment. Ditching is a premeditated maneuver

deliberately executed by the pilot with the specific intention

of abandoning the aircraft. In general, it is an act that offers reasonable hope of escape and survival. In fact, premeditated ditchings should have an equal or greater number of

survivors than forced landings on land if adequate postcrash

survival provisions are present. Analysis of the ditching environment shows stirvival can be enhanced by adequate (large,

numerous) egress openings, highly visible lighting around escape openings, and, especially for passenger-carrying helicopters, truly effective aircraft flotation devices.

6.2.1 Aircraft Configuration and Survivability

Chaeracteristics

The majority of fatalities in light fixed-wing and rotary-wing

aircraft ditchings are due to drowning. However, the behavior

of the aircraft and consequent egress difficulties vary somewhat between the two different aircraft configurations.

6.2.1.1 Fixed-Wing Aircraft: Fixed-wing aircraft generally

will remaln afloat for a sufficient length of time to permit

occupant evacuation. In a study of 306 light aircraft ditchings, Snyder and Gibbons found that, although actual flotation

times were not clear in many cases, the known data indicated

90 to 95 percent of the aircraft stayed afloat long enough for

safe egress (Reference 64). This finding is reflected in the

6T. Saczalski, K., *et al., eds., AIRCRAF1T CRASHWORTHINESS,

Charlottesville, Virginia, University Press of Virginia,

1975, pp. 121-139.

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• , L7, 1p;<'• .'.• •'m - IN, • !•4 I,,•l ':• • " '• " !
relatively high survival rates determined from the study:

88.5 percent survival for both pilots and passengers. The authors also concluded that at least 50 percent of the resulting fatalities were caused by drowning after a successful egress.

Thus, fatalities were related more often to lack of emergency

personnel flotation devices than to impact trauma or egressing

difficulties.

Aircraft configuration seems to be a factor in ditching incident survival. This same study determined that fixed-gear aircraft, whether high- or low-wing, are less successfully ditched

than retractable-gear configurations. Occupants of high-wing,

multiengine aircraft seem to have significantly less chance of

surviving a ditching than do occupants of other types of fixedwing aircraft.

6.2.1.2 Rotary-Wing Aircraft: Unplanned landings on water

are difficult for rotary-wing aircraft because their high center-of-gravity configurations are inevitably unstable in

this environment. The rotors cannot be relied upon to help keep the aircraft upright since waves may induce an early rolling tendency, causing the rotor blades to strike the water.

Also, compressor blades very often become salt-incrusted and

stall shortly after touchdown. In addition, a significant number of helicopter ditchings involve autorotation onto the

water. Flaig, in a study of U. S. Navy helicopter ditchings from 1960 through 1974, found that 24 percent of controlled,

unplanned water landings involved autorotation (Reference 65).

As with fixed-wing aircraft, most fatalities in helicopter ditchings are due to drowning. During a study of 78 Navy helicopters involved in water accidents resulting in the loss of

63 lives over a 4-year period, it was found that only 10 deaths

were due to injuries (Reference 66). Twenty-five deaths were

attributed to drownings and the remaining 28 were lost at sea.

Twenty-one, or 40 percent, of those recovered drowned or lost

at sea were last seen still in the aircraft. The overall survival rate seemed to correlate with the helicopter flotation

time, as shown in Table 16.

'65. Flaig, J. W., HELICOPTER FLOTATION AND PERSONNEL SAFETY

IN UNPLANNED WATER LANDING FROM FY 1960 THROUGH FY 1974,

Letter Report 75-61, Naval Air Systems Command, Washing- tion, D. C., June 1975.

66. Rice, E. V., and Greear, J. F. III, UNDERWATER ESCAPE FROM

HELTCOPTERS, Naval Safety Center, Norfolk, Virginial paper

presented at the llth Annual Symposium of the Survival

and Flight Equipment Association, Phoenix, Arizona, October 7-11, 1973.

146
**i'

M TABLE 16. SURVIVAL RATE VERSUS HELICOPTER FLOTATION

TIME (FROM REFERENCE 65)

Flotation Percent

time Survivors Fatalities Fatalities

<1 min 165 26 13.6

1-5 min 42 4 8.7

>5 min 83 5 5.7

In correlating fatality rates with specific helicopter models,

however, Flaig found that the helicopter flotation capability

did not correlate with the fatality rate of occupants who survived the impact but perished because the helicopter did not

stay afloat long enough (Reference 65). This seeming inconsistency results from the finding that, in larger helicopters

(more than four crew members), safety decreases faster with the

number of people than it increases with relatively good flotation. Another finding, which bears on the issue of helicopter size versus flotation, is that passengers were much more likely

to be fatalities than regular crew members. Of particular significance is the fact that 76 percent of the crew fatalities

and 92 percent of the passenger fatalities were due to drowning.

6.2.2 Underwater Escape

Since the majority of ditched helicopters roll inverted and

sink in only a few minutes, inrushing water might be expected

to hinder emergency egress. Interviews of helicopter ditching

survivors have confirmed this supposition, with inrushing water

reported as a deterrent to escape far more frequently than any

other problem (Reference 66). Inrushing water was the only

egress problem encountered by 43 survivors. However, in addition, it was reported in conjunction with several other egress

problems, as shown below:

Egress Problem Number of Survivors

Inrushing water only 43

Inrushing water plus: Reaching hatch 34

Confusion/disorientation 26

Releasing hatch 16

Darkness 12

Fire/smoke/fuel 11

Releasing restraints 9

147

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To determine the effectiveness of escape hatch illumination on

ease of egress from submerged, inverted helicopters, simulation

tests were conducted by the Naval Submarine Medical Research

Laboratory utilizing trained divers (Reference 67). Tests were

conducted using three different window escape hatches under

day and night, light and no-light conditions. The only two

variables that showed statistically significant effects were

the window used for egress and the presence or absence of window lighting. (The lighting consisted of high intensity electroluminescent lights at the tops and sides of the windows.)

The one window showing significantly longer egress times required the occupant to move the seat back support from the window in order to egress, and to exit from the window without

striking a sponson support.

More rapid egress occurred when the windows were illuminated

than when they were not. There was no significant difference

between the speed of night and day egress under either the

light or no-light conditions. Even with the use of trained

divers and controlled conditions, there were 16 recorded instances when subjects became disoriented, lost, and/or entangled within the helicopter. Fifteen of these instances occurred in the absence of illumination, and one occurred with

the lights on.

6.3 EMERGENCY EGRESS OPENINGS

6.3.1 General Provisions

Emergency escapo provisions are discussed in detail in Chapter 7. Although the provisions In Chapter 7 apply to all aircraft, the unique problems encountered in escaping a ditched

aircraft, especially rotary-wing aircraft, dictate special consideration for egress openings. Maximum egress time prior to

helicopter rollover into an inverted position and submergence

can vary from a few seconds to a few minutes. Therefore, occupant survival is highly dependent on egressing from the aircraft in a timely manner.

Since the ditching survival rate is dependent on the number of

occupants in rotary-wing aircraft, more and larger emergency

exits should be provided in passenger-carrying helicopters than

might normally be provided. The configuration of each aircraft

67. Ryack, B. L., et al., THE EFFECTIVENESS OF ESCAPE HATCH

ILLUMINATION AS AN AID TO EGRESS FROM A SUBMERGED HELICOPTER, Report No. 856, Naval Submarine Medical Research Laboratory, Groton, Connecticut, November 1977.

148
r

model dictates the potential available escape routes. Consideration should be given to providing additional escape hatches,

which can be opened if necessary, in the overhead, deck, and

tail sections to facilitate escape, especially if the aircraft

sinks on its side.

6.3.2 Explosively Cut Exits

Explosive cutting charges can be used to provide quick-opening emergency exits in downed aircraft. These systems definitely

should be considered for use in passenger-carrying helicopters

operating over water environments. Their rapid initiation time

(less than 0.1 second) and immunity to the crash environment

would provide the rapidity of opening and accessibility required of emergency exits in unplanned water landings.

Linear shaped charges should be placed around and extend beyond

existing windows and hatches to preclude the problem of jammed

or stuck exits. Strategically placed shaped charges in the

overhead, deck, and empty bulkhead spaces could provide the

additional emergency exits required in the ditching environment.

Each exit should be capable of being actuated manually and independently from the rest so that only desired exits are opened,

since opening of submerged exits may result in more rapid sinking. However, automatic actuation by water pressure could be

used after all exits are submerged.

A detailed discussion of explosively created exit systems may

be found in Section 7.2.7.

6.4 UNDERWATER EMERGENCY LIGHTING

Adequate emergency exit lighting is essential for rapid evacuation of any aircraft under conditions of reduced visibility.

It is critical in the ditching environvant because of the disorientation of aircraft occupants and the limits of underwater

visibility, even during daylight conditions. The following

sections discuss the particular problems of underwater visibility and the criteria nocossary for adequate emergency exit

lighting under water. ELergency lighting in general is discusS4 . sed in Section 7.3.

6.4.1 Underwater Visibility

The ability of an obnerver to detect an object depends not

only on the intensity of illumination but also on the visual

"threshold of the observer's eye. Smith, et al., found that

149

j ~ ~ j,, •'
luminance* thresholds in water are higher than those in ambient

air by about 1.5 log units (Reference 68). A principal reason

for this is the loss of the unprotected eye's focusing power

in the water. This loss produces severe hyparopia; that is,

the focal length of the eye is increased and the viewed target

cannot be brought into focus within the plane of the retina.

In water one does not see a sharply defined target light. but

rather a diffuse blur whose apparent size is much larger than

it would be if viewed in the air. The increase in size has

the effect of spreading the light intensity over too large an

area to be compensated for by spatial summation in the retina,

thus resulting in an increase in the luminance required for

detection.

The initial adaptation level of the eye also influences the

luminance threshold. When an observer looks from a brighter field to a dimmer field, his eyes must adapt to the change in light intensity. Thus, increased target brightness or longer

viewing times are required to compensate for the temporarily • • lower threshold sensitivity experienced when changing from

higher to lower adaptation levels. The rate of adaptation in

water parallels that in air, as illustrated in Figure 57 (Reference 68). Thus, the difference in visibility thresholds between air and water mediums is approximately 1.5 log units at

all adaptation levels.

6.4.2 Emergency Lighting Requirements

Since the curves for threshold luminance in air can be used to

predict the sensitivity of the eye in watet if the curves are

shifted downward by 1.5 log units, Smith, et al., have proposed

the following method for determining light levels necessary

for helicopter escape (Reference 68).

Bouguer's exponential law of absorption may be used to obtain

the luminance (L) required of a light source to be just visible

at a distance (d) by an individual whose threshold sensitivity

is (S) in water with attenuation coefficient (ak

*Luminance is the photometric brightness or luminous intensity

of a surface in a given direction per unit of projected area.

* It is measured quantitatively in foot-lamberts (fL) or candelas per square meter (cd/m2). One foct-lembert = 3.426

cd/ma.

68. Smith, P. F., Lurial, S. M., Ryack, B. L., LUMINANCETHRESHOLDS OF THE WATER-IMMERSED EYE, Report No. 857,

Naval Submarine Medical Research Laboratory, Groton, Connecticut, March 1978.

:1.50

• ,•.
Adapting level

+0.5 26 2 26 cd/m_ .27 cd/ m

Water

Air

• "- -0.5

U

iil ° -1.5

"4 -2.5

.,.4

-3.5

-4.5 .

2 4 8

Viewing time, sec

Figure 57. Comparison of luminance thresholds at two

adapting levels in water and in air.

L - SV (3)

where V -ead

Tho attenuation coefficient varies with climatic and water conditions. Representative values of the coefficient (a) are

.ahown in Table 17. In open water, the coefficient (a) gener- ally varies froin 0.08 to 0.125. Values in harbors, bays, and

gulfs may vary from 0.167 to 0.7, while estuaries and coastal

waters tend to be much more turbid. Conditions within ditched

helicopters may be such that the coefficient depends more on

debris or oil rather than the water in which it is ditched,

but this factor has not been evaluated.

f 151

____ __L
TABLE 17. REPRESENTATIVE ATTENUATION COEFFICIENTS

(a) FOR VARIOUS WATER SOURCES

Water source a Year of determination

Pacific Countercurrent 0.083 1951

Pacific North Equatorial Current 0.083 1951

Gulf of Mexico (Panama City) 0.100 1967

Pacific South Equatorial Current 0.111 1951

Caribbean Sea 0.125 1951

Caribbean Sea (Roosevelt Roads) 0.300 1969

Long Island Sound 0.700 1967

Thames River (Connecticut) 3.500 1969 IA

Viewing distance (d) will vary with seating arrangements and

escape hatch placement.

Sensitivity (S) will vary among the aircraft crew and flight conditions. Occupants not looking outside the aircraft may be

exposed to adapting fields of 15 to 50 cd/ma. Pilots and crew

members who must look outside the aircraft will be exposed to

much higher levels. For example, a pilot flying in a hazy sky

can experience 25,000 to 35,000 cd/mA. During ditching, pilots

will be looking at the water, which is generally less bright than the sky, and adaptation levels will be reduced to approxi- ¶ 4mately 350 cd/ma.

The minimum output levels (IL) for escape hatch lights may be '.J

determined by substituting appropriate values of V and S in

Equation (2). The value of V can be found from Figure 58,

which gives values of V for various attenuation coefficients

and the distance from the light to the observer. Figure 59

provides threshc-ld sensitivity levels for elapsed times following extinction of an adapting field with luminances from 0.35

to 350 cd/mr.

60i 152

l~ 1 
....... i

100,000 a-10.0 a=3.5 a=2.5 a-2.0 a=1.5

1,0,000

a=1. 0

1,000

a=~0.25

2100 

a10 a=0.25

0 1 2 3 4 5 6 1 8 9

Distance, meters

Figure 58. Values of V in Equation (3) for distances from

2ight source to observer with various water attenuation coefficients (a). (From R~eference 68)

153

. . . . . . . . . . . . TO .
i•¢ •,•, ... . .,.l

0

>4

".4

; : •AF a 35

wA a

P 20

_i5}:. o 2 5 1o02

i•!' Time, sec.

' •-•1;•Figure 59. Threshold sensitivity (S) levels for elapsed

:/ time after ditching for adaption fields (AF)

-:•, from 0.35 to 350 cd/ . (From Reference 68)

15

!"2

"• .'•',,, .h',~~~~~~~~~A ,,a," .. 3 50, : •;',, ' '';,' : '•

', , ,;,,,, ,_ _ __,,_',_ __,_ __-'_ __'__"'__. ,AF, - ..

"IL '.... "" '"AF...
A degree of latitude is available in selecting the sensitivity

level. Several seconds are required during a ditching for

conditions to stabilize enough to permit escape. During this

time, ambient light levels are dropping, thus lowering the

adaptation levels. Requiring the lights to be visible 2 sec

after ditching, rather than immediately, significantly lowers

the threshold sensitivity (S), as may be seen in Figure 59.

Consequently, the lower value of S will reduce the required

brightness of the exit lights.

Table 18 lists the values of L obtained from Equation (3) when

the adapting level is 350 cd/m2 and the light must be visible

immediately upon ditching. The effect of increasing turbidity (higher attenuation coefficient) on light level requirements

is apparent. The calculated luminance (L) is the minimum value

at which the light is just barely visible. The higher the

value that L is above the threshold level, the greater the

probability of detection. Therefore# the highest brightness

level of light permitted by other design conditions should be

employed for the escape lighting system. Based on the data in

Table 18 and the sizes of passenger-carrying U. S. Army helicopters, the minimum brightness of the lights should be at

least 120 fL (411 cd/M2 ).

TABLE 18. EXAMPLES OF VALUES FOR MINIMUM LUMINANCE (L) IN

cd/m2 FOR ESCAPE HATCH LIGHTS AT DISTANCES TO

LEVEL OF 350 cd/mr (From Reference 68)

Distance Attenuation coefficients

(meters) 2.5 1.0 0.5

0.5 34 17 14

1.0 127 27 17

1.5 439 45 21

2.0 1,538 75 27

2.5 5,331 127 34

3.0 18,610 206 45

3.5 64,950 339 58

4.0 * 562 75

5.0 - 1,528 127

6.0 - 4,152 206

7.0 - 11,289 339

8.0 - 30,680 562

9.0 83,399 925

*Values below here become very large and are perhaps prohibitive.

155

',

S. . ... ..- , •, •
6.5 AIRCRAFT FLOTATION SYSTEMS

Several methods currently being used in attempts to provide

ditched helicopters with flotation capabilities include inflatable bags, large sponsons, sealed hulls, and combinations

thereof. Some of these methods have not been particularly successful in preventing postcrash fatalities, since they were

unable to provide adequate flotation times for the escape of

all occupants from larger helicopters. For instance, although

one type of Navy helicopter has floated upright for more than

2 min in 70 percent of its ditchings, it has a high fatality

rate (Reference 65).

If large numbers of people are to be carried, the flotation

provisions must be very effective to lower the fatality rate.

As might be expected, the number of inadequate flotation incidents will decrease as more flotation provisions are incorporated in any given helicopter. Thus, consideration should be

given to using a combination of flotation methods, such as

sponsons in conjunction with flotation bags, sealed hulls, etc.

6.5.1 Sponsons

Although sponsons are not intended to permit extended periods

of operation on water, they can help stabilize the aircraft in

relatively calm seas. However, to be of any value in providing

flotation, the sponsons must be quite large to counteract the

inherent instability due to a helicopter's high center of grav- ity.

The sponson buoyancy required to stabilize an aircraft for

small angles of rotation may be estimated by using the following equation (Reference 69):

tan 0 - (4)

* where 6 - heeling angle, deg

F B = maximum single sponson buoyance, lb

e = horizontal distance from aircraft centerline

to the center of buoyancy of the sponson, ft

d - vertical distance of the aircraft center of P : buoyancy to the aircraft center of gravity, ft

W = normal gross weight of the aircraft, lb

156

. . I .. 
The heeling angle calculated for Equation (4) should be verified by data from tests performed on the aircraft or on a scale

"* model thereof.

6.5.2 Flotation Bags

Inflatable gas bag flotation systems have been developed and 4:..

are currently being used on several aircraft. Their success

to date, however, has been limited, Reliability problems have

yet to be solved satisfactorily (Reference 65). In addition,

buoyancy requirements of truly effective flotation bags pose

design problems relative to the size and location of the deployed bags.

The flotation bag buoyancy required to stabilize a helicopter

to any desired heeling angle may be estimated from the following equation (Reference 69):

t Fe + Pbeb tan e - d W - (5)

where 0 w heeling angle, deg

Fs a maximum single sponson buoyancy, lb

es - horizontal distance from aircraft centerline to

center of buoyancy of the sponson, ft

Fb a buoyancy of inflated bag, lb

eb a horizontal distance from aircraft centerline to

center of bag, ft

d - vertical distance of aircraft center of buoyancy

to aircraft center of gravity, ft

W - normal gross weight of aircraft, lb

As may be seen, the maximum heeling angle determined from EquaV tion (5) is dependent on the buoyant force of the bag (bag

2' size) and the distance of the bag from the aircraft.

4, 69. Saczalski, K., et al., eds., AIRCRAFT CRASHWORTHINESS,

Charlottesville, Virginia, University Press of Virginia,

1975, pp. 645-667.

157

iN

4 1<' 4i4j
To auhieve maximum effectiveness, the bags should be inflated

simultaneously, prior to or at water contact at a low airspeed.

Reliability considerations of the flotation system are of prime

importance. The failure of both bags to inflate, or the separation of both bags from the aircraft upon water contact,

will destroy any effectiveness the system might have. Moreover,

the loss of buoyancy on one side could cause the aircraft to

list and possibly sink faster than it would without the system.

6.6 DITCHING EQUIPMENT

Suitable tiedown or stowage facilities should be provided for

life rafts, life preservers, survival kits, and miscellaneous

ditching equipment. Restraint devices and supporting structures for equipment should be designed to restrain the equipment to static loads of 50 G downward, 10 G upward, 35 G forward, 15 G aftward, and 25 G sideward. All survival equipment

should be readily available and easily released from their restraining devices by the occupants after ditching. More details on the design requirements for containing emergency

equipment may be found in Volume 3 under Aircraft Ancillary

Equipment.

Provisions for carrying life rafts should be included in all

aircraft whose mission requires frequent flight over water,

especially if the aircraft mission also includes troop transport. Research has shown that individuals are not able to tolerate exposure in 320F (0OC) water for more than 90 min, or

500F (100C) water temperatures for more than 18 hr (Reference

70). Figure 60 shows that a life raft between the sea and the

individual provides a significant buffer that extends the tolerance time for a period of days. A raft with an effective

spray canopy can make the difference in survival of aircraft

occupants in the sea.

The design and location of life raft mountings or restraining

devices should be such that rafts can be removed from their

mounts or enclosures and deployed outside the aircraft within

30 sec from the time that release or removal action is initiated by the operator.

When exterior installations for life rafts or other survival

equipment are provided, the mountings, retention devices, or

enclosures should be designed to preclude inadvertent actuation

or damage in flight or when ditching. Such equipment should be

70. Veghte, J. D., COLD SEA SURVIVAL, Aerospace Medicine, Vol. 43, No. 5, May 1972.

158

. .. . . . . . . . . .. . ... • Z c • •• • -. .. .U
35 -Shipwrecked

- Laboratory S30- 4 t least one survived c' All survived

~25-

S20 I

S10

1. 260 14 16225 30 34 3842 460 426 H4562 6670 74

Time, hr

Figure 60. Tolerance time for different water

S temperatures while in life rafts.

recoverable by occupants from an exit intended for use in ditching. Release mechanisms should minimize the possibility of jam- ming due to structural deformation that might be incurred upon

ditching.

(... 159

a :14
7. EMERGENCY ESCAPE PROVISIONS

7.1 INTRODUCTION

Even though an occupant has survived a crash, the problem of

surviving the postcrash environment still remains. Nonsurvivable postcrash conditions occur in a relatively small percentage of accidents, but they account for a disproportionately

lazge number of injuries and fatalities. The key to postcrash survival is the time between the initial crash sequence and

the onset of nontolerable conditions. The primary postcrash

hazards are fire and water. The occurrence of either can reduce the available escape time to seconds. Therefore, effective emergency escape provisions are essential as integral por- tions of the aircraft design.

7.2 EMERGENCY EXITS

7.2.1 Emergency Exit Requiremants

Two factors that largely determine emergency exit requirements

are (1) the amount of available time before the postcrash environment exceeds human tolerance limits, and (2) the attitude

and condition of the aircraft structure after it comes to rest.

Research has shown that the available escape time from helicopters involved in postcrash fires is only 7 to 16 sec (see

Chapter 3). Thus, all occupants must be able to evacuate the

aircraft within 10 sec if they are to survive. However, the

allowable evacuation time can be extended to 30 sec if a crash-,

resistant fuel system is installed in the aircraft. The emergency exit criteria presented in this chapter are predicated

on the installation of sudh a fuel system and should allow all

occupants of an aircraft to evacuate within 30 sec.

Providing sufficient exits for 30-sec evacuation of the maximum

number of personnel to i•e carried would seem to meet the emergency requirements. However, it is not unusual for several

exits to be blocked following a crash. For instance, if a

rotary-wing aircraft comes to rest on its side, all exits on

"that side will be unusable. Also, exits can be blocked by

outside objects, such as trees, or by deformation of the airL I craft structure. Therefore, emergency escape provisions should ,? allow the maximum number of aircraft personnel to evacuate in

30 sec with only one-half of the aircraft exits available for

Evacuation times should be demonstrated by actual tests using

personnel approximating 95th-percentile troops with full. combat

160

terA
•" equipment for passengers and 95th-percentile aviators with arctic flight gear and body armor for crew members. (Anthropometric data for U. S. Army aviators can be found in Volume 2.)

The following sections present emergency exit design criteria

to assist the designer in meeting the above requirements.

7.2.2 Types of Exits

Aircraft exits are provided to facilitate either normal or

emergency exit and entry. Generally, these exits are classified as follows:

9" Class A Exit: A door, hatch, canopy, or other

exit intended primarily for normal entry and exit.

e Class B Exit: A door, hatch, or other exit intended

primarily ?or service or logistic purposes (e.g.,

cargo hatches and rear loading ramps or clamshell

doors).

e Class C Exit: A window, door, hatch, or other exit

intended primarily for emergency evacuation. Exit closures for Class C exits must be capable of being

removed from the exit opening within 5 sec regardless

of the aircraft's attitude.

A Class C exit constitutes the minimum requirement for an emergency exit. A Class A exit with emergency jettison provisions j is normally considered superior to a Class C exit because of

its large size, and, in most cases, it can be used in lieu of a

Class C exit. Despite its superiority, however, each Class A

exit with emergency jettison provisions can replace only one

Class C exit. Class B exits also may be used in lieu of Class

C exits if adequate emergency release provisions are installed;

however, the functional design of Class B exits usually makes

their use legs desirable for emergency exit. In order for

Class A and B exits to qualify for use in lieu of Class C

exits, the exit closures must be capable of being removed from

the exit opening within 5 sec regardless of the aircraft's attitude.

7.2.3 Size of Exits

"All exits, including Class C exits, must be large enough to

accommodate 95th-percentile troops and aviators as specified

in Section 7.2.1. Furthermore, the exits must be large enough

to allow these personnel to evacuate the aircraft rapidly.

Class C exits should be at least 22 in, square with 6-in-.radius

corners, or 22 in. in diameter. This exit size is an Air Force

161

iA,
requirement and is considered to be minimum for the evacuation

of troops at the rate of 1.5 sec per person (Reference 71).

This size must be considered an absolute minimum since the anthropometric data given in Volume 2 lists the shoulder breadth

of a 95th-percentile U. S. Army aviator as 20.3 in. Therefore,

it is strongly recommended that all Class C exits be larger

than the minimum 22 in. Other shapes may be used also, providing the minimum dimensions are met or exceeded. In any case,

all exits must be sufficient in size and shape to allow 95thpercentile troops and aviators, equipped as specified in Section 7.2.1, to pass through the exit at a rate of 1.5 sec per

person or less.

7.2*4 Number of Exits

7.2.4.1 Crew Compartment (Cockpit), Each flight crew member

must have access to at least one usable emergency exit regardless of aircraft attitude after impact. Thus, if a single

cockpit enclosure is used for a single crew position, two

Class C exits on opposite sides of the cockpit should be provided. This arrangement assures an alternate means of escape

if the aircraft rolls on its side, blocking one exit. One

Class A exit with an emergency release provision may be substituted for each Class C exit if desired. The Class A exit

may be the normal entry/exit door with an emergency jettison capability.

The minimum emergency escape exit requirement for cockpit enclosures with side-by-side crew positions is also two Class C

exits. One exit should be installed on each side of the fuselage. Although two Class C exits are required, any comb:Lnation

of Class C and Class A exits may be substituted, provideid the

Class A exits have an emergency jettison feature.

Cockpit enclosures with tandem crew positions should be provided with two Class C exits and one Class A exit with an emergency release provision. This requirement assumes that the

two crew positions are mutually accessible. Mutual accessibility means that a 95th-percentile crew member dressed in arctic

flight gear and body armor could, without undue difficulty,

climb from one crew position to the other in order to exit the

aircraft in an emergency. If the exits in such a cockpit are

not mutually accessible to the crew members because of intervening structure or installed equipment, each crew member

should be provided with a Class C axit and a Class A exit with

71. Air Force Systems Command, DESIGN NOTE 3Q6; in AFSC:

DHL-6, System Safety, 4th Edition, Wright-Patterso--Air

Force Base, Ohio, July 1974.

162
I. .... • • t•• • •: • , •• .. . .

Aan emergency release provision. When sliding or clamshell canopies are used, Class C exits or other suitable means should

be provided for crew escape when the aircraft is inverted or

otherwise malpositioned on the ground and tha canopy cannot be

jettisoned. Accident records for aircraft with canopy-type

cockpit enclosures indicate crew members are often trapped in

the cockpit when the aircraft flips over on its back and the

canopy cannot be jettisoned. Experience with such accidents

indicates that knives, axes, or other tools carried in the • cockpit for chopping through the Plexiglas canopy are not adequate solutions for emergency exit when postcrash fire or occupant injury is present. When the primary means of escape is

blocked, an alternative means is clearly necessary.

7.2.4,2 Passenger or Troop Compartments: The minimum emergency escape exit requirement for troop/passenger sections,

exclusive of exits provided in cockpit sections, is two Class

C exits. One exit should be installed on each side of the

fuselage. If one of the two exits becomes blocked for any reason, the other exit will serve as the primary means of escape.

Normally, a Class A exit is required for passenger/troop compartments. If normal passenger or troop entry and exit in a

particular aircraft is through the troop/passenger compartments, a Class A exit with emergency release provisions and a

Class C exit will be more realistic and satisfy the emergency

exit requirements.

In addition to the minimum number of exits, additional exits

may be required depending on the maximum number of personnel

carried in the pas:;enger/troop cabin. Class C exits at a ratio of at least oii( exit for every 10 persons expected to occupy the section 0Iould be provided. An additional exit in

excess of tne 1-to-10 ratio should be provided when the specified capicity of the section is not evenly divisible by 10

(e.g., if the capacity is 21, three exits' are required). The

requirement for the passenger/exit ratio of 1-to-10 is based

upon the possibility that at least 50 percent of the exits may

be blocked if the aircraft comes to rest on its side. This

1 would then leave a 20-to-i ratio, assuming that both sides of

the aircraft have an equal number of exits. A 20-to-l ratio is,

from a theoretical point of view, considered adequate to evacuate all occupants within 30 sec at the exit rate of 1.5 sec per

person (assuming no troop debilitation and all exits open).

However, at least two exits must be provided even if the number

of occupants is less than 10.

The exit requirements cited above also are applicable to cargo

compartments if the compartments have a dual capability for

troop transport.

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t 3
7.2.5 Location of Exits

7.2.5.1 Side Exits: Exits inteaded for emergency use should

be equally divided on eich side of the aircraft and, if feasible, should not be directly opposite each other. The primary

reason for dividing emergency exits equally on both sides of

the fuselage is that an alternate means of escape is provided

if, for any reason, the exits on one side bccome blocked.

Exits should not be located directly opposite each other because of the probability of crowding in one particular arer,

woen both sides of the aircraft may be used for evacuation.

By staggeiing the exits, the tendency to crowd up is dimiaiahed.

Since any aircraft may be operated over water, at least one

emergency exit on each side of the fuselage should be well

above the anticipated waterline under the most adverse conditions expected after a water landing. 

'

7.2.5.2 Overhead Exits: In aircraft where the width of the

crew and troop compartments is too great to permit easy access

to fuselage up-side exits if the aircraft comes to rest on the

opposite side following an accident, Class C overhead exits

should be provided at a ratio of one exit for every 20 occupants. Where the capacity of the compartment is less than 20,

at least one Class C exit should be present. These overhead

exits are in addition to the normal requirements for Class 

C

exits.

When an aircraft comes to rest on its side, blocking the exits

on that side, the exits on the other side of the aircraft could

be the only means of evacuation. These exits, now on the topsid- of the rolled aircraft, may be useless if the width of

StY. Zuselage is such that they cannot be revi 3d easily. In r•. an aircraft resting on its side, overhead •i,. would be more

accessible than the normal up-side exits. A fuselage width of

5 ft or more between side exits is considered too great to permit easy access to up-side exits by troops with minor debilitating injuries following a crash.

In helicopters with engines, transmissions, major controls,

etc., located over personnel compartments, bottom or fore

and/or aft exits may be substituted for the overhead exits.

Alternatively, side exits may be located where interior aircraft components, such as seats and consoles, can be used as

steps to gain access to thn up-side exits. If this type of arrangement is used, the designer must ensure that these components will maintain their structural integrity and attachment

to the aircraft during a survivable crash. Such component-

&teps must be able to support a 300-lb occupant to accommodate

fully equipped 95th.-percentilu crew members and troops.

164

&.
If the aircraft has a higl-wing arrangement, overhead exits

T should be provided to facilitate escape following ditching. These overhead exits will be in addition to the normal requirements for emergency exits. Overhead exits constitute the

only practical means of escape in a rapidly sinking aircraft

of this type because the occupiable portion of the fuselage in

high-wing aircraft sinks below the surface of -he water rapidly

following a ditching. The opening of side exits causes flooding of the interior at a high rate, decreasing escape time.

7.2.5.3 Exit Location Relative to Fuselage Distortion: To

provide maximum accessibility to aircraft occupants following

a crash, emergency exits should be located in areas least

vulnerable to distortion. Insofar as it is feasible, exits

should not be located in close proximity to the main landing

gear because of the possibility of the gear being driven upward and/or inward against the aircraft, causing a blocked or

jammed exit. Exits should not be located under heavy components mounted on the top of the fuselage, such as engines and

transmissions, because of the possibility of fuselage distortion in crashes where high vertical forces are present. In

high-wing aircraft, a crash landing is likely to cause structural deformation below the wing; therefore, exits located under the wing should be avoided as much as possible.

7.2.5.4 Exit Location Relative to Obstructions: Class C exits

should be located where it will not be necessary to move equip- r. nment, cargo, or furnishings to gain access to them. Insofar

as it is feasible, all exits that might be used in emergencies

should be located where external components such as propellers,

turbine engine inlets, turrets, armament, and tail surfaces

will not interfere with occupant escape.

7.2.5,5 Exit Locations Relative to Ignition Sources: Exits

should be located as far as possible from fuel spillage areas

and from major ignition sources (e.g., exhaust stacks, hot engine parts). Where the occupiable portion of the aircraft is

mainly aft of Lhe power units and fuel tanks, it is desirable

to locate at least one Class A or B exit with an emergency jettison featuxre as far aft as possible. In the case of rearmounted engix'es, an A- or fl-type exit should be as far forward

as possible. Such an arrangement may increase escape time in

the event of a postcrash fire.

7.2.6 Exit Operation

7.2.6.1 Exit Operational Design: The method of emergency exit

operation should be simple, obvious, and natural to all personnel expected to be aboard the aircraft. Exit operation also

should be as rapid as possible. Therefore, exits intended for

165

;kl 1".

i i 

i
emergency use should be designed so that no secondary operation such as moving or unlocking locks, catches, stops, bolts, or

bars is necessary. (Such a requirement does not preclude the

use of easily removable protective covers intended to prevent inadvertent actuation of exit release handles.)

An emergency exit should be capable of being completely opened

within 5 sec. The time requirement of 5 sec to remove the exit

closure (window, door, hatch, etc.) from its opening is based

upon the need for all possible haste in evacuating burning aircraft and a realistic estimate of the time-motion requirements

for actuating a simple, continuous-motion release mechanism

without secondary operations. The measurement of time should

begin when the operator places his hand on the release handle

and end when the exit closure is free and clear of the exit

opening.

Only the single operation of pulling or pushing the exit closure into the clear should be necessary, once the release handle has been actuated. Unless the aircraft is pressuzized,

all emergency exit closures should be arranged to fall free,

or: to be easily pushed outward tor side exits, to be pushed

upward for overhead exits, to be pushed downward 5or bottom

exits, when the emergency release mechanism is actuated. To

remove the exit closure inward would add to the congestion and

impede escape. In a pressurized aircraft, exit closures must

be removed inwardly, but, if at all possible, the closure

should then be manted at an angle and pushed out the exit opening in order to avoid congestion inside the aircraft.

Emergency exits should be designed to permit removal of the

exit closure when seal vulcanization occursi when the fuselage

is covered with ice accumulated in flight, anid when minor fuselage deformation occurs. A peripheral clearance of at least

0.20 in, provided between the exit closure and its frame will

help accomplish this goal.

The 0.20 in. specified should be considered the minimum clearance between the exit closore and its frame. It is probable

that some aircraft with relatively light fuselage construction

*i could use more than 0.20-in. clearance in this area, since

greater fuselage distortion in such aircraft is likely when a

crash occurs. With a 0.20-in. peripheral clearance, the exit

frame could theoretically deform inward for 0.40 in. on any

one of its four sides before binding occkrs.

Consideration also should be given to designs that cause the

exit closure to eject itself from its frame when large structural deformation due to impact occurs. This type of design

is particularly appropriate for the simple Class C type of exit

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mI.

,i * J
that contains no release mechanism but needs only to be pushed

out of its mountings to open.

The use of Class A and B exits that slide to open is probably

unavoidable on certain types of aircraft. Such a design should

not be incorporated on exits intended for emergency evacuation 

,

because fuselage distortion, which is common in aircraft

crashes, can cause the exit to bind on the tracks attached to

the fuselage.

7.2.6.2 Release Mechanism Design: The exit release mechanism

is the primary handle, lever, or latch used to open the emergency exit closure. Handles may be of the T- or L-shaped design that turns, the D-ring type that pulls, or the lever, type

that slides fore and aft. However, the number of different

types of handles in the aircraft should be held to a minimum.

It is recognized that some types of emergency exits will not

use exit release handles. One common type of exit uses a release method whereby a panel held by a flexible mounting is

simply pushed out.

The method of operation of the exit release mechanism should

be simpli, obvious, and natural to the operator. In order to

facilitate rapid emergency egress, exit release mech:anis;,:i

should be designed to permit release handle actuaton and exit

opening by one person using one hand. The Air Force specifies

an actuation/operating force of 10 to 30 lb to meet this requirement (Reference 71). Release and opening mechanisms also

should allow all exits to be removed successfully in an emergency when the aircraft is in other than an upright position,,

The shape and direction of operation of exit release handles

should conform to the "form follows function" rul-e, where the

releasing action is most natural to the position of the operator initiating the action. According to McFadden and SweariA ingen, "In general, the best position for applying force to a

handle is one in which a subject can use his legs and lift.

The next best is in pushing down and using body weight. The

least effective method is the employment of an over or under

motion. The under motion is slightly superior." (Reference

72).

Specific considerations for different types of handles are as

follows:

72. McFadden and Swearingen, FORCES THAT MAY BE EXERTED BY MAN

IN THE OPERATION OF AIRCRAFT DOOR HANDLES, Journal of the

Human Fc. tors Society of America, Vol. 1, No. 1, September

17958.

., 167
T- or L-Shaped Emergency Release Handles: Internal

emergency release handles with a T- or an L-shaped

design should be capable of actuating the release

mechanism in both clockwise and counterclockwise directions. The arc of rotation in this case should

not exceed 90 degrees. If only one direction of handle rotation is permitted, rotation should be counterclockwise and ths arc of rotation should not exceed 180 degrees. Stops that prevent rotation in

the wrong direction should be provided.

9 D-Ring Type Emergency Release Handles: If the release handle is a D-ring type that requires pulling for release action, the grip of the D-shaped handle

should be parallel to the aircraft's vertical axis

for side exits and parallel to the aircraft's longitudinal axis for overhead exits. The direction

of pull should be toward the operator in the same

straight line as the natural position of the extended forearm holding the handle prior to release

action.

e Lever-Type Emergency Release Handles: Internal emergency release handles incorporating a lever or bar

that slides fore and aft along the x axis of the aircraft should be capable of opening the exit in both

directions.

Exit release mechanisms should be designed so that the entire

operation of the release handle is a continuous motion from

start to finish without sharp changes in direction except for

external installations where the release handle must be pulled

from countersunk recesses before actuation. In any type of

release handle, the final motion of the handle should contribute to the opening of the exit.

Release handle shapes and dimensions should be designed for

normal hand grip limitations and incorporate handle-to-hand

contact areas that ensure adequate load applications to the

handle. Release handles on external installations should provide clearance to allow gripping of the handle with gloved

hands, since rescue crews normally wear heavy gloves to protect hands frowl jagged and hot metal surfaces. Standard fire

fighter's asbestos gloves should be used for testing. The release handle should be mounted on the exit closure itself or

immediately adjacent to the exit opening so that it is readily accessible to any occupant attempting to use that exit.

Remote exit release mechanisms should be avoided. The release

168

a aNi
rL 

r

handles on the exit closure or on the adjacent airframe should

not be located in a position that would allow the handle to

snag clothing or impede escape through the exit opening even

if the exit is, for some reason, limited to partial opening. Similarly, the exit actuating mechanism should be designed so that the final position of the release handle upon opening will

not obstruct the removal of the exit closure.

Emergency exit release handles in cockpits and troop compart- * ments should be located where it is not necessary for crew members to unlock their shoulder harnesses in order to turn the

handles. This is very important in cockpits and at crew chief

or special crew stations, primarily because it is sometimes

desirable to release emergency exits just prior to crash impacts. This is especially true for ditching. If a shoulder

harness has to be unlocked to release the exit, there may be

insufficient time available to relock it before impact. This

requirement also is applicable to those emergency exits that

are adjacent to certain seats in the passenger/troop compartment. This, however, should not be construed as a recommendation to romove exits prior to crash impact in every case. The

openings of such exits can sometimes critically reduce the time

otherwise available for occupants to escape, since fire can

develop on the outside, causing flash fires within the compartment. As a general rule, the chances of surviving a crash involving fire are less if doors and exits are open prior to im- 

- pact.

Accidental release of exits in flight can be extremely dangerous in rotary-wing aircraft. Exits that have been released in

flight have been known to fly into the main or tail rotor sys- tem, causing disintegration of the system and subsequent loss

of the aircraft and crew. An unguarded or unshielded exit release handle can make a convenient hand-hold for inexperienced

troops. Therefore, release mechanisms should be designed so

that improper or incomplete closing of the exit will be obvious.

On both external and internal installations, a locked-position indicator, such as a detent to indicate positive locking, should be provided.

In the event that crash victims become trapped in the aircraft

or become otherwise unable to escape without help, it is essential that all emergency exits be capable of being opened by

rescue personnel from the outside of the aircraft. The actuation of an internal release handle must not preclude the simultaneous actuation of an external release handle. If "push-out"

type Class C exits are provided, they should be as easy to open from the outside as from the inside. Means to prevent icing of

the outside release and handle mountings should be provided to

ensure positive operation under adverse weather conditions.

169

!4 I~ 

y4
If the cockpit enclosure consists of a canopy that slides back

and forth or opens on hinges in a clamshell fashion, an emergency jettison feature can provide rapid egress for the crew.

The jettison mechanism should allow complete removal of the

enclosure from its mounting within 5 sec from the time that

mechanical action is initiated. In addition to the internal

jettison release, external canopy jettison controls should be

provided on both sides of the fuselage. The canopy jettison feature does not eliminate the necessity for additional emergency exits since the postcrash aircraft attitude might pre- clude successful jettisoning of the canopy.

7.2.7 Explosively Created Exits

Explosive systems have been developed and successfully used to

provide quick-opening emergency uxits in military aircraft.

These systems can cut emergency exits through existing doors

and windows and through fuselage structures. The systems provide the advantages of extremely rapid release times, simplic- ity of operation, and immunity to jamming by structural deformation, ice, or seal vulcanization. The following sections

discuss factors that must be considered during the design of

an effective and operationally safe explosive exit system.

7.2.7.1 Overall System Design: An explosively operated exit

system contains four basic components or subsystems: (1) an

arming/firing system, (2) primer and/or detonating cord, (3) a linear shaped cutting charge, and (4) an actuation mechanism. The relationship of these components to each other can best be

illustrated by considering the design of an actual system--in

this case, the Emergency Lifesaving Instant Exit System (ELSIE)

developed for the U. S. Air Force (Reference 73).

The ELSIE system is composed of an electromechanical safe/arm

mechanism, dual shielded mild detonating cord lines, a flexible

linear shaped cutting charge, and interior and exterior initiation handles attached to firing lanyards. The relationship of

the components is shown schematically in Figure 61. The safe/

arm mechanism requires only momentary application of power to

arm or disarm. The system remains armed or disarmed, even if

power is lost, since the mechanism is mechanically locked in

position. Once the system has been armed, it can be actuated

73. Nicholson, D. E., and Burkdoll, F. B., AN EMERGENCY LIFESAVING INSTANT EXIT SYSTEM FOR CARGO, CARGO-TRANSPORT AND

PASSENGER AIRCRAFT, Explosive Technology; ASD Technical

Report 71-41, Aeronautical Systems Division, Air Force

"Systems Command, Wright-Patterson Air Force Base, Ohio,

June 1971, AD 736056.

170

.. - . .. .T,}
Initiation

Shielded mild manifold

detonating cord

il; ~~Exterior ,• .

' ~handle

initiation;

Flexible

linear "shaped charge

interior \A

Safe/arm initiation handle

mechanism Jettisonable panel

Safe/arm i

switch Energy-absorbent Detail A-A

... Linear J material shaped charge / •

Moisture barrier

Aircraft skin

lFigure 61. Schematic of ELSIE system. , . (From Reference 73)

either from inside or outside the aircraft by pulling the handle in either location. The handles operate a mechanical striker that fires the dual detonating cord lines. These redundant lines, in turn, initiate the shaped charge that cuts

the egress opening in the aircraft and ejects the cut panel

outward. Tests on the ELSIE system show that the elapsed time

from pulling the initiation handle until the egress opening is available for use is less than 0.027 sec.

171

.. ... ... ...
7.2.7.2 Arming/Firing System: The arming/firing system must

be designed for simple and rapid actuation of the explosive

system and yet provide maximum safety against inadvertent eLctuation. Operational safety must be assured by preventing inadvertent actuation due to environmental conditions, system

component failures, or human error.

To provide maximum operational safety, arming and firing must

be accomplished in two separate and deliberate actions. The

arming function always should be under the control of the

flight crew. Thus, the arming mechanism should be located

only in the cockpit and at the crew chief's station. If cockpit enclosures with tandem crew positions are used, each crew

member must be provided with an arming mechanism unless the

two positions are mutually accessible. System status indicators should be provided at all pertinent flight crew stations.

Once armed, the system should be capable of being fired by any

of the aircraft occupants. Each exit should be capable of

being actuated independently from the rest since it is not always desirable to open all available emergency exits, especially in case of a postcrash fire or a ditching. A firing

mechanism should, therefore, be located immediately adjacent

to each exit for actuation of that particular exit only. This

means the arming and firing mechanisms will, of necessity, be

physically separated from each other. An exception to this

practice might be acceptable when the exits are located quite

near each other, as in tandem cockpit configurations. Then

the adjacent exits could be fired simultaneously from one

firing mechanism, although a firing mechanism must still be

available to each crew member.

Once the system is armed, it must stay armed until it is disarmed by u crew member or rescuer. The reverse also is true;

once the arming mechanism is in a disarm, or safe, position,

it must remain that way until a deliberate arming action is

initiated. Any type of system or component failure must not

change the position of the safe/arm mechanism. For instance,

if arming is accomplished by electrical power, loss of power

must not allow the mechanism to switch from arm to safe or vice

versa. The mechanism also must be immune to any environmental

or crash load input. Disarming capability must be provided

to permit safing the system even though normal safing modes

are inoperable following a crash.

In order to provide the highest degree of both operational and

crash safety, the firing mechanism should be independent of

any external energy source, such as Lhe aircraft electrical

system. This requirement dictates that the firing mechanism

be manually operated. The design considerations for emergency

172

Ifr (
exit release mechanisms discussed in Section 7.2.6.2 also apply

A :• to the firing handles used in explosive exit systems. In addition to those considerations, the external release handle must

be designed to allow rescue personnel sufficient separation

from the aircraft before actuation to prevent their being

struck by debris when the exit is opened. It is also strongly recommended that all arming mechanisms and firing handles be

completely separated from each other, even in those cases where

it might seem feasible to combine them (e.g., pilot's crew station). If the arming and firing mechanisms are combined into

one package, it is essential that the operations of arming and

firing be distinctly separate from each other, such as turning

the handle to arm and pulling the handle to fire.

7.2.7.3 Explosive System: All explosives used in the exit

system should possess as high a thermal limit as possible, not

only to ensure that the system is safe in high-temperature op- erating environments but also to provide as much safety as pos- sible in case of a postcrash fire. The system must be able to

function when exposed to temperatures up to the limits of human

tolerance to heat (approximately 400*F, based on ambient air

temperature), yet not function inadvertently during brief exposure (30-60 sec) to postcrash fires. The latter requirement is necessary to prevent flames coming through an unintentionally opened exit of an occupied aircraft. The thermal limits

of the explosives used in the ELSIE system, which meet the

above requirements, are below.

Explosive Thermal limit (_OF)

lHNS (22', 44', 66' hexanitrostilbene) 618

SLead azide 635

II M-426 primer 425

The linear shaped charge must be held securely in position "against the aircraft structure it is to cut. The size of the

exit opening should conform to Class C requirements given in

Section 7.2.3. The jettisonable section must be ejected outward to preclude its obstructing the exit opening. Energy- absorbing backup material must be placed behind the shaped

charge to control the backblast of the explosive and prevent fragments from entering the cockpit or cabin (refer to Figure 61).

"The explosive system must be designed to minimize the possibility of system actuation igniting any fuel that might be spilled

during a crash. The amount and duration of any exposed flame

173

J_ _ _ .... __, _ __. _ _U " "
must be minimal. The ELSIE system successfully functioned during a series of fuel spray tests without igniting the fuel because the explosive charge was designed to penetrate only 90

percent of the aircraft skin thickness. The remaining 10 percent was severed by the pressure created by the detonation of

the shaped charge and the momentum already imparted to the jettisonable section. This design allowed the combustion products

around the periphery of the cut to cool significantly before

the metal skin was completely severed. Because of this, the

only flames exterior to the aircraft skin were at the initiation points of the shaped charge and lasted less than 10 msec

for most of the tests.

7.2.8 Access to Exits

7.2.8.1 Exit Obstructions: Access from aisles to all exits

must be provided so that the exits will not be obstructed by

troop seat components, seat back webbing and webbing support

bars, litter installations, or other protrusions to an extent

that would reduce the effectiveness of the exit.

A common problem with troop-carrying aircraft is that, in order

to carry the maximum number of troops, some emergency exits

are blocked by the installation of troop seat back webbing and

webbing support bars. These components are normally designed

to be pulled away from the emergency exit in order to provide

access. It is desirable, of course, to avoid obstructing exits,

but if it is necessary to do so, seat backs or other potential

obstructions should be readily collapsible or movable to provide access to exits during an emergency evacuation.

7.2.8.2 Aisle Widths. The width of aisles at any point between seat rows should be sufficient to allow unobstructed

movement of 95th-percentile troops with full combat equipment.

Current criteria suggest a minimum width of 17 in. In aircraft

where it is necessary to pass through seat rows to gain access

to exits during an emergency, longitudinal spacing between seat

ruws should be sufficient to permit these troops to move at a

rate consistent with the capacity of the exit (1.5 sec per person or less).

7.2.8.3 Compartment Doors: Doors on hatches separating various interior compartments should be located and have a direction of opening so as not to impede or block passage to other

exits or interfere with the use of emergency equipment such as

axes and fire extinguishers.

The doors or hatches should be large enough to permit crew and

troop movement from one compartment to another during emergency

evacuation. The openings should have no protrusions that would

impede movement through them. Provisions should be made for

174

i ,,;
securing compartment doors in the open position during takeoff

and landing. Such doors should be capable of remaining open and latched when exposed to crash forces of survivable magnitude.

Compartment doors should have release handles designed so that

the method of operation is a single, obvious, and natural mo- Stion in a single plane. Round or spherical handles or knobs

should unlatch when gripped and turned in either direction.

The handles should not snag on clothing or equipment.

7.3 EMERGENCY LIGHTI1G

Emergency lighting provides the illumination required for emergency evacuation and rescue when the normal aircraft lighting

is not available. There are three basic types of emergency

lighting: (1) interior lighting for personnel orientation following aircraft accidents at night, (2) lighting for the purpose of reading exit operating instructions and releasing the

exits, and (3) exterior lighting to illuminate exits and paths

of escape.

7.3.1 Interior Emergency Lighting

When an aircraft crashes at night or is filled with dust or

smoke, disorientation of embarked crew and 'troops is likely to

occur. Since escape time is critical, interior emergency

lighting units should be installed in sufficient number and

possess adequate brightness to permit personnel orientation in

all compartments during emergency evacuation situations. The

emergency lighting should provide sufficient illumination

throughout the cockpit and cabin areas to permit occupants to

patsan locate emergency avoid exits obstacles and while survival moigtoward equipment, perceive the exits. escape This

criteria may not he necessary for some aircraft using overhead

canopies.

Interior lighting fixtures may be mounted as aisle, ceiling,

or cornice lights, Regardless of where the lights are mounted,

* they must furnish adequate illumination near floor level to

allow occupants to see exit paths and avoid any obstructions.

The emergency lighting requirement for both civil and military aircraft is a minimum average illumination in clear air of 0.05

foot-candle (fc) measured 20 in. above the floor (or at armrest

height) along passageways leading to each exit and in front of

A175
each exit (References 74 and 75). (1 foot-candle is the in- tensity of illumination 1 ft from a source of 1 candela.)

Full moonlight is approximately 0.04 fc, while recommended reading intensities are 20 to 50 fc (Reference 76).

7.3.2 Emergency Exit Lights

Supplementary emergency lighting units should be provided at

or near each emergency exit with adequate brightness to permit untrained personnel to identify exits, to read exit operating instructions, and to actuate the exit mechanism without difficulty during periods of reduced visibility. The identity and

location of each emergency exit should be recognizable under

limited visibility (darkness, smoke, etc.) from a distance

equal to the width of the cabin.

Exit light requirements must take into account the fact that

the illumination at any distance from a light source is inversely proportional to the square of the distance from the

source. Thus, at a distance of 5 ft, the brightness of a

light will theoretically diminish to only 4 percent of the

brightness measured 1 ft from the source. The same rapid decrease has been measured in the brightness of internally illuminated aircraft exit signs during an FAA program to evaluate

current exit signs and markers (Reference 77). The decrease

in average exit sign brightness with increasing distance from

the signs is shown graphically in Figure 62.

Exit light effectiveness also is reduced by the presence of

smoke. Measured light output for all units tested by the FAA

diminished proportionately in a 90 percent smoke environment,

as shown in Figure 62.

74. Federal Aviation Regulations, AIRWORTHINESS STANDARDS:

TRANSPORT CATEGORY ROTORCRAFT, Part 29, Section 29.811.

75. Military Specification, MIL-L-6503, LIGHTING EQUIPMENT, AIRCRAFT, GENERAL SPECIFICATION FOR INSTALLATION OF, Department of Defense, Washington, D. C., 2 April 1975.

76. Mendenhall, C. E., et al., COLLEGE PHYSICS, 3rd. Edition,

Boston, Massachusetts, D. C. Heath and Company, 1950,

p. 476.

77. Garner, J. D., and Lowrey, D. L.,,EXIT SIGN COMPARISONS

IN CLEAR AIR AND SMOKE, unpublished paper presented at an

A-20 meeting, Society of Automotive Engineers, New York, December 1976.

176 3
100

90

No smoke

80 90% smoke

70

S60

50\

S40

350

20

30

20 *\

10i0 I- - I• 

1 2 3 4 5

e Distance from sign, ft

Figure 62. Relative brightness of internally illuminated

aircraft exit signs at various distances from

signs (average of 10 different signs).

Current FAA requirements for large transport-category airplane

emergency lighting include internally or self-illuminated signs

at each exit with a minimum luminance (brightness) of at least

25 fL (Reference 78). Small (9 seats or less) transport airplanes and transport-category rotorcraft need only have exit

signs with a brightness of 160 microlamberts (0.15 fL) (References 74 and 78). Although the above requirements might be

78. Federal Aviation Regulations, AIRWORTHINESS STANDARDS:

TRANSPORT CATEGORY AIRPLANES, Part 25, Section 25.811.

"177

!I J

• ' : .. . ' ,: ' .'•I I '
I, I

sufficient in cleat air, the rapid drop in brightness due to

%he presence of smo;,e makes the sufficiency of even the brighter (25 fL) requirement questionable.

Most current transport airplane exit lights exceed the 25 fL

requirement, but lights far brighter than those currently usedI

are available. Figure 63 presents the results for two of the

10 exit lights tested by the FAA (Reference 77). This figure shows that, under all conditions, the experimental light was approximately 10 times brighter than the typical currently used

exit light. This is most important during smoke conditions

and at some distance away from the exit sign. For instance, at a 6-ft distance under 90 percent smoke, the current sign

transmitted only 0.017 fc of light while the experimental sign transmitted 0.13 fc. It is noteworthy that the experimental 9

aircraft sign is currently used in building installations and

uses less battery power than some, current aircraft signs.

Other newly developed lights, which are much brighter than current lights, also are available.

Based on the results of the FAA tests, all passenger- or troopcarrying aircraft should contain internally illuminated exit

signs with a minimum average brightness of at least 25 fL.

However, it is strongly recommended that the exit signs be even

brighter.

The diminishing of exit light effectiveness when the aircraft

is submerged has already been discussed in Chapter 6. Any aircraft whose mission requirements include troop transport over

water should contain exit sign lighting meeting the requirements specified in Chapter 6.

7.3.3 Exterior Emergency Lighting

For noncombat missions, exterior emergency lighting should be

considered at each exit to illuminate the ground near the exit

and areas where escape and survival equipment will be deployed. MIL-L-65C3 specifies that the light intensity on the ground below normal and emergency exits should be 0.02 fc minimum

(Reference 75).

7.3.4 Structural Considerations

All emergency lighting units must be self-contained, explosionproof, oparable under water, and accessible for periodic maintenance. All units must be capable of operating independently of the main aircraft lighting system.

The emergency lighting system should be designed, installed,

and located so as to minimize damage to or loss of any portion

178

~-'~.-
16 * - Smoke (90%)

i .... No smoke

14 Experimantal sign (407 fL)

Current sign * (55 fL)

12

4J

6$

4

~:' I

2

0 !.2 3 4 5 6 ~

Distance from s~zgn, ft

Figure 63. Light transmittance of exit signs in clear air

arnd in smoke at various distances from sign.

179

4-$ I4$
of the emergency illumination as a result of ditching or emergency landing. To ensure structural integrity and continued K

operation after a crash, the lighting system, including all

components necessary to provide the reqlired illumination,

should be capable of withstanding the following crash loads:

50 G downward, 10 G upward, 35 G forward, 15 G aftward, 25 G

lateral. Breakup of the fuselage should not render any portion of the emergency illumination inoperative except for those

lights directly destroyed by the break.

7.3.5 Emergency Lighting Power Source

Emergency l.ghting power sources should be independent of the

main electrical power source for the aircraft and should contain power sufficient to ensure effective illumination for a

minimum of 15 min.

It is believed that a power source strong enough to provide at

least 15 min of effective illumination following a crash at

night is adequate. If a postcrash fire does not occur within

15 min, it is likely that one will not occur at all. Personnel

who are stunned or otherwise unable to evacuate the aircraft

during the 15 min of emergency lighting could, in all probability, evacuate in the darkness if they were physically able.

7.3.6 Actuation of Emergency Lighting Units

Emergency lighting units should be designed to actuate both

automatically and manually. If inadvertent actuation occurs,

the unit should be capable of being reset manually.

7.3.6.1 Manual Actuation: There age circumstances where it

would be desirable to manually turn on the emergency lighting.

One such instance would be when a crash was imminent, but some

time was available prior to the crash. By turning on the emergency lights manually, the aircraft occupants would have time

for their eyes to adjust from normal lighting or darkness to

emergency lighting. Thir also would permit all normal aircraft

lighting to be tu.ý ied off in order to reduce potential postcrash fire ignition sources. Therefore, a manual actuating

switch should be placed in the cockpit, and another should be

placed in the passenger/troop compartment close to the crew

chief's station.

7.3.6.2 Automatic Actuation: The emergency lighting units

should be automatically actuated in as many survivable accidents as possible. This can be accomplished by using inertia

sensors responsive to the crash pulse parameters typical of

lower-severity accidents. The sensor criteria should be identical to those specified for crash locator beacons (see Chapter 8).

180
The crash sensors may be contained in each lighting unit, or

the units may be actuated from one or more common sensors lo- S, , cated remotely from the lights. The circuits for the lights

should be such that they will be energized if the circuits between the lights and the sensors are broken.

There may be circumstances, such as forced or crash landings

in enemy territory, where it would not be desirable to automatically actuate the emergency lighting. A circuit breaker

or other device to nullify the automatic feature therefore is

desirable. Such a device could be utilized by the crew upon entering enemy territory.

7.4 MARKING OF EMERGENCY EXITS

Emergency exits must be clearly marked both inside and outside

the aircraft so that occupants and rescuers can find them rapidly. The markings must be distinctive to set them apart from

the numerous other markings found on the aircraft. In addition

to identifying the exits, instructions for releasing the exit

closures must be clearly marked beside the exit release mechanism. The time required to determine how to release the exit

closure could well mean the difference between survival or nonsurvival.

All U. S. Army aircraft must be painted and marked according

to the requirements in TB 746-93-2 (Reference 79). The requirements contained therein for marking of emergency exits

are summarized in the following sectionu. The reader is referred to TB 746-93-2 for complete dotails.

7.4.1 Internal Identification of Exits

An orange-yellow band must mark the complete periphery of the

escape exit on olive drab backgrounds. A gloss black band is

used on light backgrounds. The band must be between I and 2

* 'in. wide and divided equally, if possible and practicable, between the mounting of the exit and the exit itself.

If soundproofing (or lining) covers the identification band

on the inside of the aircraft, it also must be appropriately

marked.

The words EMERGENCY EXIT, in orange-yellow, must be marked or

stenciled on the escape hatch, door, or exit in the most readily visible location. Preferably, letters should be 2 in. high,

but they cannot be less than 1 in. high.

79. Technical Bulletin 746-93-2, PAINTING AND MARKING OF AIUIY

AIRCRAFT, Department of the Army, Washington, D. C., January 1971.

181
7.4.2 External Identification of Exits

Markings identifying escape hatches, doors, and exits on the

outside of aircraft must be marked gloss yellow on dark surfaces and gloss black on light surfaces. On camouflaged aircraft, emergency exit markings are painted with lusterless yel- low lacquer.

7.4.3 Identification of Secondary Openings

Secondary openings, such as auxiliary exits and windows, are

usually smaller than primary openings, making entrance or exit

more difficult. If the structure immediately surrounding a

secondary opening is free from heavy structural members; oxygen, fuel, and oil lines; and battery leads it should be

marked with a broken orange-yellow band on olive drab or a

broken black band on light Surfaces. The band must be placed

on the extreme boundary of the clear area around the secondary

opening, inside and outside of the fuselage. Segments of the

broken band must be 1/2 in. wide, 1 in. long, and approximately

12 in. apart. CUT HERE FOR EMERGENCY RESCUE must be marked

inside of, parallel with, and adjacent to the broken band identifying the area on the outside of the aircraft where forced

entry can be made for rescue purposes. CUT HERE FOR EMERGENCY

EXIT must be marked in a similar location inside the aircraft.

Letters must be I in. high.

On camouflaged aircraft, the corners of emergency exits and

rescue exit areas are outlined with right-angle corner bands

1 in. wide and 3 in. long at each leg. The corner markings

are painted with lusterless yellow lacquer.

7.4.4 Marking Instructions for Exit Operations

7.4.4 1 Internal Markings: Small handles or levers used to

actuati doors or hatches must be identified by alternate 1/8-

inch-wide orange-yellow and black stripes, painted on the background of the exit. Background striping must be applied at a

15-degree angle from the vertical, rotated clockwise. The

striping must not interfere with other types of markings or

codings. Large levers or exit controls must be marked with

alternate orange-yellow and black stripes, 1/8 to 1/4 inch

wide, painted directly on the lever or control.

7.4.4.2 External Markings: All external releases for operation of energency exits must be marked EXIT RELEASE on the outside of the aircraft to facilitate quick identification. Letters preferably should be 2 in. high and must not be less than

i in. high.

182
7.4.4.3 Operating Instructions Markings: Operating instruct tions to identify and explain the emergency release operation

must be marked on the exit door, or hatch, or aircraft structure, whichever is nearer the release. Minimum lettering

heights specified are 1/2 in. internally and I in. externally.

Preferably, the descriptive wording should be 1 in. high on the

inside of the aircraft and 2 in. high on the outside. The

1/2-in. minimum specified in TB 746-93-2 for internal wording is not sufficient for easy reading under reduced visibility conditions, such as darkness or the presence of smoke.

The instructions should be as simple and concise as possible

consistent with clarity of meaning. Standard English terminology, such as PULL, PUSH, TURN, or SLIDE, must be used.

The painting and marking schemes for in-service aircraft contained in TB 746-93-2 show liberal use of nonverbal symbols in

exit operating instructions. Symbols are particularly useful

in delineating directions of motion for handles, levers, etc.

The use of symbols in conjunction with words will often lead

to quicker understanding of the operation to be performed.

Symbols are invaluable when the wording cannot be deciphered,

as might be the case under reduced visibility conditions, or

when viewed by non-English-speaking personnel. Thus, although

not stated as a specific requirement in TB 746-932, symbols

should be used in exit operating instructions whenever possible. Some symbols in current use are shown in Figure 64.

7.5 CREW CHIEF STATIONS

There is a great need for experienced crew chief personnel to

provide the necessary leadership and guidance for embarked

troops during emergency evacuation. Accident records indicate

that on many occasions crew chiefs have been responsible for

successful emergency evacuations of large numbers of troops from aircraft under severe conditions.

At least one crew chief station should be located in each troop "compartment. The location of the crew chief's station should

provide as complete surveillance of the troop compartment as

is practicable. The station should be located as near the main

or emergency exits as possible. For aircraft requiring two

crew chiefs, their respective stations should be as far apart

as practicable; e.g., one in the forward end of the compartment

and one in the aft end.

183

* -\
0 0()000

0

0

0 OPEN~

0

0s00

0 00

EXIT RELEASE

EXIT REL~EASE

TURN

Figure 64. Typical exit release instructions

incorporating symbols.

184
8. CRASH LOCATOR BEACONS

8.1 INTRODUCTION

After a survivable crash has occurred, rescue time becomes

paramount in determining the ultimate survival chances of the

occupants. Air Force records indicate that the life expectancy

of injured survivors decreases as much as 80 percent during

the first 24 hr following an accident, and the chances of survival of uninjured occupants rapidly diminishes after the first

three days (Reference 80). Therefore, the installation of a

crash locator beacon in the aircraft can greatly enhance the

occupant's survival chances by reducing the amount of time between crash and rescue. The following sections present criteria that should be followed to ensure the satisfactory operation of a crash locator beacon installed in an aircraft.

8.2 TYPES OF LOCATOR BEACONS

Crash locator beacons can be either automatically or manually

activated, dnd they can be permanently installed (fixed), portable, or automatically deployed.

8.2.1 Automatic and Manual Activation

Automatic activation of the transmitter is triggered when the

crash sensor detects a preset impact condition. Such activa-

* tion requires no previous action on the part of the crew and,

for that reason, is the preferred method. However, in military

aircraft there are obviously times when it is not desirable to

activate a locator beacon on a downed aircraft. For this reason, an arming switch should be included to provide the option of automatic activation depending on the aircraft mission. A

manual activation switch also should be provided so that the

crew can activate the beacon after the crash if the arming

switch is not on or if, for any other reason, the beacon is

not automatically activated.

8.2.2 Fixed, Portable, and Deployable Equipment

Fixed equipment is permanently mounted in the aircraft. The

transmitter, antenna, knd power supply need not be contained

in one package, althou,_ +-heir close proximity to each other

lessens the chasnes foý. coz,.icting circuitry to be damaged in

the impact. Portable L.nd Jý-Ioyable beacons, on the other hand,

must include the transmitter, antt4,ina, and power supply in one

package.

80. EMERGENCY LOCATOR TRANSMITTRRSt AN OVERVIEW, Report No.

NTSB-AAS-78-1, National Transpoz':ation Safety Board, Washington, D. C., 26 January 1978.

185

; -I)'7t
Portable equipment is designed to be easily removed from its

installation in the aircraft by crew members for use in a location remote from the downed aircraft if so desired. However,

its installation in the aircraft must be secure enough to protect it from impact damage. In addition, an externally mounted

antenna should be provided so that the beacon can function in

the desired manner if it is not removed from the aircraft.

Automatic deployable equipment is automatically ejected from

the aircraft when the crash sensor experiences crash forces

equal to or greater than the present values. The equipment

must be designed to minimize or withstand ground impact forces

after ejection. Also, it should be buoyant, self-righting,

and stable when floating in water, and not adversely affected

by immersion in fresh or salt water for the life of the power

supply. Automatically activated and deployable equipment

should be considered for any aircraft that spends a considerable portion of its operating time over water.

8.3 COMPONENTS OF LOCATOR BEACONS

All aircraft-installed crash locator beacons contain the same

basic components: a crash sensor, transmitter, antenna(s),

power supply, activating switch, and associated electrical circuitry.

8.3.1 Crash Sensors

8.3.1.1 Types of Crash Sensors: Several general types of sensors can be used to detect aircraft crashes. These include:

9 Ground contact switches that sense abnormal landings

in which the fuselage or a wing tip is in contact

with the ground.

* Proximity switches that sense the altitude of the

d. 'aircraft and determine if the aircraft is below

the normal level at which it would be supported on

its landing gear.

* Deformation or damage indication switches that sense

deformations of landing gear or wings and/or displacement of engines.

* Inertial switches that sense the decelerative loads

applied to the aircraft.

* Attitude switches that sense sustained odd attitudes

typical of aircraft wreckage after a crash.

186

* • •1

!• " •, , ,,• - •/"!• -••' t" '
e Aircraft system sensors that detect excessive pressure drops in fluid systems.

9 Tachometer-operated switches that detect excessive

changes in the speed of rotor blades.

If inertia switches are used, they must be capable of integrat- ing acceleration over time so they can differentiate between a 4 crash situation and normal aircraft operational accelerations.

*• Two such sensors are currently available for aircraft. One is

based on a mass held in place by a magnet, and the other is

based on a mass resisted by a spring. These sensors are described in References 32 and 81.

Currently, the majority of crash locator beacons use inertia

sensors to activate the transmitter. The FAA has specified

crash force criteria for transmitter activation, although alternate means may be used if they can be shown to be substantially equivalent in their response to the specified crash

forces (Reference 82). Regardless of the type of sensor used,

the sensor must be responsive to the majority of survivable

aircraft accidents, including those accidents in which the

crash forces and damages are minimal. In this latter regard,

an inertia sensor is probably more universally applicable than

a sensor that responds to only local conditions in various

parts of the aircraft.

8.3.1.2 Criteria for Inertia Crash Sensors: Inertia sensors

must be designed to respond to the aircrafE crash loads but to

ignore normal operational loads. These two environments must

be defined before the sensor criteria can be determined.

8.3.1.2.1 Crash Environment: A detailed analysis of over 500

survivable Army aircraft accidents has resulted in the definition of crash pulses for both light fixed-wing and rotary-wing

aircraft. (This effort is discussed in Volume II.) The analy- sis provided frequency of occurrence data for average fuselage floor accelerations and aircraft velocity changes experienced during the crashes.

81. Johnson, N., and Sanderson, S., SPILLED FUEL IGNITION

SOURCES AND COUNTERMEASURES, Ultrasystems, Inc., Dynamic

Science; Report No. DOT-HS-801 722, U. S. Department of

Transportation, National Highway Traffic Safety Administration, Washington, D. C., September 1975.

82. Federal Aviation Technical Order, EMERGENCY LOCATOR TRANSMITTERS, (TSO) C91.

187

aim=
The median average longitudinal acceleration was only 2.7 G,

with 50 percent of the accidents experiencing lower accelerations. Since the crash locator beacon should activate in as

many survivable accidents as possible, any G-type sensor must

respond to acceleration levels as low as 2 G to capture 75 to

80 percent of the accidents. However, combinations of vibrational loads and flight loads can sometimes exceed this acceleration value. Therefore, some other parameter, such as velocity change, also must be specified to prevent inadvertant

sensor actuation.

The median longitudinal velocity change determined for the aircraft crashes was 28 ft/sec, with 95 percent of the accidents

experiencing a velocity change of 13 ft/sec or more (see Figure 65). Thus a G-sensor could be designed to filter out the

vibration and flight environment, even with a low 2 G threshold

level, if it also must detect a velocity change typical of crash

rather than operational conditions.

Fixed-wing aircraft rarely have sizable vertical crash forces

without also experiencing large longitudinal forces. Helicopters, on the other hand, can experience large vertical crash

loads with minimal longitudinal loading. Average vertical accelerations were determined during the aircraft accident anal- S~ysis mentioned previously. The median vertical acceleration

was 3.2 G, slightly higher than that of the longitudinal accelerations. However, the average vertical acceleration for 80

percent of the accidents was 2 G or more, the same as for the

longitudinal accelerations.

The median vertical velocity change was 24 ft/sec, with 95 percent of the accidents experiencing a velocity change of 10 ft/

sec or more (see Figure 66).

The aircraft accident survey produced insufficient data to

yield an accurate distribution of lateral crash accelerations

or velocity changes. However, significant lateral accelera- tions were present in the crashes studied, particularly in some

types of helicopter accidents. The data did suggest that the

maximum lateral accelerations and velocity changes were somewhat less than those along the longitudinal and vertical axes.

8.3.1.2.2 Operational Environment: Operating conditions for

fixed-wing aircraft can be estimated by using the vibration

test loads specified by the Radio Technical Commission for .... iAeronautics (RTCA) in their performance standards for emer-

* gency locator transmitters (Reference 83). These loads are:

* 83. MINIMUM PERFORMANCE STANDARDS - EMERGENCY LOCATOR TRANSMITTERS, Document No. DO-147, Radio Technical Commission

for Aeronautics, Washington, D. C., 5 November 1970.

188
60

50

'44

40

U

30 Median - 28 ft/sec

0

" 20

.0 2: 2 0 4 0 60 7 0 -0 9 0

.4.

10

0

0 10 20 30 40 50 60 70 80 90 100

Cumulative frequency of occurrence, percent

Figure 65. Longitudinal velocity changes for survivable

rotary- and light fixed-wing aircraft accidents.

i I

189 7

, -N
50

o 40

4J4.

V 30

U

S ~Median- 24 ft/sec ,

0

r 20 .....

0)

S1 0... .

0 10 20 30 40 50 60 70 80 90 100

Cumulative frequency of occurrence, percent

Figure 66. Vertical velocity changes for survivable rotaryand light fixed-wing aircraft accidents.

* Constant total excursion of 0.100 in. from 5 Hz to

that frequency where a peak-to-peak acceleration of

10 G is reached. '-

. From the frequency determined above (31.28 Hz) to

2,000 Hz, a constant peak-to-peak acceleration of

10 G.

The total velocity changes under the half-sine waves corresponding to the above vibration spectrum are listed in Table 19.

Although peak accelerations are greater than the desired 2 G

..90

,' A-], 9
AI

TABLE 19. VIBRATION SPECTRUM CHARACTERISTICS

BASED ON RADIO TECHNICAL COMMISSION

FOR AERONAUTICS, DO-147 REQUIREMENTS

Frequency Acceleration Double amplitude Velocity change

(Hz) (±G) (in.) (ft/sec) ¶

5 0.13 0.10 0.262

10 0.51 0.10 0.524

20 2.07 0.10 1.047

30 4.62 0.10 1.571 I

31.28 5.00 0.10 1.633

40 5.00 0.061 1.279

50 5.00 0.039 1.038

100 5.00 0.001 0.513

threshold, the velocity changes due to the vibration spectrum

are considerably less than those experienced during 95 percent of the aircraft crashes.

However, as pointed out by the Crash Research Institute, the

above treatment is too simplistic because the vibration environment is actually superimposed on the flight environment

(Reference 84). In developing crash sensor criteria for fixedwing aircraft, the Institute assumed a possible 1.5 G longitudinal flight force (which could occur under the unlikely conditions of a flight load limit of 6 G and pitch attitudes of ±15

degrees of the flight path) and superimposed it on top of the

worst vibration load listed in Table 19. The results are depicted graphically in Figure 67. The velocity change under

[* these conditions (the shaded area under the curve) was found

to be 2.7 ft/sec, still considerably les4,than accident velocity changes.

84. DEVELOPMENT OF ELT CRASH SENSOR PERFORMANCE SPECIFICATION

AND TEST PROCEDURES - FINAL REPORT, Crash Research Institute, Robertson Research, Inc.1 Contract No. DOT-FA76WA3842, Department of Transportation, Federal Aviation Administration, Washington, D. C., June 1977.

191

O''

-----
5 G pkpk 31.28 Hz

Worst vibration load

• +a

2 G

A' 0

-- 3.5 G a - Longitudinal

0 Time . L component of

severe flight load, 1.5 G

Figure 67. Worst vibration loads superimposed on severe

longitudinal components of flight loads.

This same type of procedure can be used to estimate operating

conditions for helicopters. The vibration spectrum used is

that specified in MIL-STD-810B for equipment installed in aircraft and helicopters (Reference 85). The vibration test

curve used (curve Z of Figure 415-1, Reference 85) is more severe than any of the other curves with the exception of those

for equipment installed in the engine compartment, or pylon,

or on the engine. The vibration spectrum characteristics for

this curve are listed in Table 20. Although there are two inflection points in the MIL-STD-810B vibration curve as compared to only one in the RTCS curve, the velocity changes for

each of the two curves are in the same general range.

Another parameter to be considered in aircraft performance is

the vertical acceleration experienced by the aircraft during

operational conditions. Vertical flight loads for helicopters

can be significant and must be taken into account in determining the operational environment of rotary-wing aircraft. Assuming a 2.5 G vertical flight load (specified for the Army's

Advanced Attack Helicopter) on top of the worst vibration loads

of Table 20, the maximum vertical velocity change is 3.0 ft/sec.

The maximum velocity change using the vibration curve of Table

19 is very similar -- 2.9 ft/sec.

8.3.1.2.3 Criteria: Aircraft accident data have shown that

an inertia sensor must have an actuation threshold as low as

2 G acceleration in order to detect 80 percent of all aircraft crashes. However, the analysis of the previous section

85. Military Standard, MIL-STD-810B, ENVIRONMENTAL TEST METHODS,

Department of Defense, Washington, D. C., 11 June 1967.

192

i i+ A*li i iIi+i+ i .. , -? + -+liii? : • + + ,+ ,,, , " ' ' '. . ; .. .. + +Iiri[ . + ,my. +' C:• ' . .+ I I
TABLE 20. VIBRATION SPECTRUM CHARACTERISTICS

* .BASED ON MIL-STD-810B REQUIREMENTS

Frequency Acceleration Double amplitude Velocity change

(Hz) ___ (in.) (ft/sec)

5 0.13 0.10 0.266

10 0.51 0.10 0.522

20 2.00 0.10 1.024

30 2.00 0.043 0.582

33 2.00 0.036 0.621

40 2.94 0.036 0.742

50 4.60 0.036 0.942

74 10.00 0.036 1.383

80 10.00 0.031 1.280

100 10.00 0.020 1.024

500 10.00 0.001 0.205

has shown that vibration and flight loads during normal operations can exceed this limit. The specification of d minimum

velocity change that must be detected by the sensor before

it will actuate, in addition to a threshold G level, can prevent actuation during normal conditions while assuring actuation during crash conditions.

The C-ash Research Institute has recommended a threshold acceleration limit of 2 G and a minimum velocity change of 3 ft/

sec for crash sensors in light fixed-wing aircraft (Reference 83). These specification limits are shown in Figure 68.

The velocity change of 3.0 ft/sec is above the maximum 2.7 ft/

sec longitudinal velocity change considered typical of the

light fixed-wing operational environment. This value also

should be satifactory for helicopters, which experience even

lower maximum longitudinal flight loads. The 3.0 ft/sec velocity change is only marginally satisfactory in the vertical

direction, though, since expected vertical velocity changes

193

A,,

~ n•
10 -

AV 3.0+_3 f t/ sec

S 8 -

0

.4.

o 6-

Sensor must activate

'4J

9-t

0

Gth =2 0+.2

S" - 0

ý4 2

> Sensor must not

activate

0.

0 20 40 60 80 100

Pulse duration, msec

I•Figure 68. Proposed specification for crash force

sensors in light fixed-wing aircraft.

for helicopter flight operations can be expected to equal or

slightly exceed this limit. The problem is compounded by the

inclusion of energy-absorbing landing gear that will allow a

normal landing at sink speeds up to 10 ft/sec.

Since 95 percent of the aircraft accidents experienced vertical

velocity changes of 10 ft/sec or more, the sensor criteria can

reasonably be modified in the vertical direction to a minimum

10 ft/sec velocity change with a 2 G threshold level. The specification limits for this case are shown in Figure 69. In

194
40 AV 10.+1.0 ft/sec

0

-W 30

rd

Sensor must

4) activate

U

S20

(U

S10

SSensor must G 2.0 + 0.2

not activate _ th 0.. 

1,1 0 I I .I I I I .

0 20 40 60 80 100 120 140 160

Pulse duration, msec

Figure 69. Proposed specification for vertical crash

force sensors in rotary-wing aircraft.

fact, any desired velocity change can be combined with the 2 G

threshold level to generate a family of curves similar to those

in Figures 68 and 69. However, to prevent inadvertent actuation during flight maneuvers, the specified velocity change should not be below 3 ft/sec. Any velocity change above 10

ft/sec will reduce the number of accidents that will be sensed.

(A velocity change of 15 ft/sec will sense approximately 85

percent of the accidents.)

Another parameter that must be considered is whether the crash

sensor should be unidirectional, bidirectional, or omnidirectional. Inasmuch as most fixed-wing aircraft accidents have a

major longitudinal component of velocity and force, a unidirectional sensor mounted with the active axis forward in the direction of the longitudinal axis of the aircraft should be sufficient. The situation is different for helicopters, which

may have large vertical crash forces with minimal longitudinal

forces. Thus, a vertically oriented crash sensor, as well as

a longitudinal crash sensor, should be used in helicopters. As more data become available, lateral crash sensors might be

indicated also.

195 [ "

. , 3) 
The sensor must be able to withstand the impact forces associated with severe survivable crashes and still function. Thus,

the sensor must withstand shock pulses equal to or greater than

those listed in Table 21.

TABLE 21. SUMMARY OF DESIGN CONDITIONS

FOR CRASH LOCATOR BEACONS

Velocity Peak

Impact change acceleration

direction (ft/sec), (G)

Forward 50 35

Downward 42 50

Lateral 30 25

Upward 10

Aftward - 15

8.3.1.3 Criteria for Other Sensors: There are no generally

applicable criteria for other types of crash sensors at the

present time. Accident data to determine optimum sensor locations or deformations in the majority of survivable crashes are

not available. However, because of the vibratory and installation problems that can occur with inertia sensors, the Crash

Research Institute is currently working with NASA to develop a

crash locator beacon that may use an "odd attitude-long exposure" type of crash sensor. This type of switch sensen3 abnormal attitudes that persist for relatively long periods of time

(e.g., 1 min). As yet, however, no criteria are available for

this type of sensor.

8.3.1.4 Sensor Mountng: The inertia sensor criteria recommended in the preceding section are based on crash forces present in survivable crashes. These are the forces seen at the

aircraft floor and thus are typical of the forces transmitted

to the occupant compartment. Therefore, the crash sensor must

be located in an area that will experience crash forces representative of those that will be seen in the occupant compartment. Of course, the sensor must be protected from any impact

damage that could render it useless before it is able to activate the transmitter.

196

A

--... .. ' ' . ' . 1 1 /,AA†
The sensor must be mounted to rigid structure, not flexible

sheet metal bulkheads that may amplify or attenuate flight or

crash loads. Amplification could lead to inadvertent actuation, while attenuation could result in a less severe crash

being missed. For the same reasons, the crash sensor must be

rigidly mounted to the structure. Soft mounting techniques,

such as flexible straps or Velcro fasteners, must not be used.

If portable beacons are installed, and if the sensor is integral with the rest of the equipment, the beacon must be rigidly

mounted to the structure. Alternatively, the sensor can be

rigidly mounted remotely from the rest of the equipment and a

quick-disconnect provided for the connecting circuitry. This

would allow easy-access installation of the portable equipment.

8.3.2 Transmitter

8.3.2.1 Mounting: The transmitter must be protected from any

impact damage that could render it useless. Therefore, the

transmitter must be located in an area that is not subject to

impact damage. Generally, this would be above the aircraft

floor and near the center or rear of the fuselage.

The transmitter must be able to withstand the impact forces of

a severe survivable accident and still function as designed.

This means, in addition to retaining its component and structural integrity, the transmitter's mounting to the aircraft ; must meet the same static attachment strengths as those speci- fied for all ancillary equipment. These are:

9 Downwards 50 G

N Upward: 10 G

e Forward: 35 G

e Aftward: 15 G

e Sideward: 25 G

8.3.2.2 Activation: The transmitter should be capable of

being either manually or automatically activated. An arming

switch that will allow the automatic activation capability to

be selected or not, as desired, should be provided. Manual

activation must always be available in case the sensor malfunctions or unusually low-level accelerations fail to trigger the

sensor. The cockpit should be provided with a warning light

that could alert the crew to inadvertent transmitter activation.

A manual override switch should be provided so that the transmitter can be turned off whenever desired.

197

...... ....... ..- i.. ' .......... ,,........... ..... ,I' : . .. ... ....-......-... 
8.3.2.3 Operating Characteristics: Operating frequencies and

transmitter ranges (power) must be determined by the procuring activity according to its own special needs. The FAA has spe- cified a 75 mW peak effective radiated power since this would

provide a transmission range of 50 nautical miles under unfavorable crash site conditions (Reference 82).

8.3.3 Antenna

Antennas, except for those used in portable and automatic deployable equipment, are usually mounted outside the aircraft

to ensure the proper radiated signal strength and shape. Since

survival of the antenna is critical to the successful operation of the crash beacon, care must be taken in lecid. r..r its location. It must be kept out of primary impact v•nes, such as

the front or bottom of the aircraft, and it also shoulO be kept

out of secondary impact zones. These zones include wing or taii surfaces likely to impact trees, etc., and those portions of helicopters apt to experience rotor blade strikes during

impact. The strength of the antenna attachment also must be

sufficient to withstand decelerative impact forces.

8.3.4 Power Supply

The crash locator beacon must have its own independent power supply so that it is not dependent on aircraft power for its operation. The power supply should be capable of providing

necessary power for optimum transmitter operation over a spe- cified time period and under specified environmental conditions.

These conditions should be specified by the procuring activity

dependent on particular mission requirements. The FAA, in following the RTCA requirements, has specified that the power sup- ply must be able to provide continuous operation for at least

48 hr under the condition of maximum power consumption (Reference 83). Maximum low and high operating temperatures are

specified as -20 0C (-40F) and 550C (131*F), respectively. However, the National Transportation Safety Board (NTSB) has

pointed out that the low temperature standard is not sufficient

to ensure operation of the crash locator beacon during winter

* •in many areas of the United States, and recommends a low temperature standard of -400C (Reference 80).

The power supply, if not integral with the transmitter, must

be mounted to the aircraft so that it will not be torn loose

or damaged during impact. It should be miunted in a location

away from anticipated impact areas and s.,ould have an attachment strength equal to that specified for the transmitter.

198

19
8.3.5 Electrical Wiring ii

All electrical wiring between components of the system must be

protected from impact damage unless the components are packaged together. Protection can be accomplished by routing the wires

along the strongest structural members of the aircraft and away

from anticipated areas o... structural deformation. The wires

should be attached to the aircraft structure with clamps or

ties that will fail before the wires break. Twenty to 30 percent extra length in the wires, in the form of loops or Sshaped patterns, will allow the wires to move with deforming

structure rather than be pulled apart. Nonconductive shields

should surround the wires in all areas where structural crushing could occur.

199

LILL.",
REFERENCES

S1. Military Standard, MIL-STD-1290(AV), LIGHT FIXED- AND

ROTARY-WING AIRCRAFT CRASHWORTHINESS, Department of Defense, Washington, D. C., 25 January 1972.

2. Johnson, N. B., et al., AN APPRAISAL OF THE POSTCRASH

FIRE ENVIRONMENT, Dynamic Science (The AvSER Facility);

USANLABS Technical Report 70-22-CE, U. S. Army Natick

Laboratories, Natick, Massachusetts, September 1969,

AD 699826.

3. CRC-Aviation Fuel Safety Report Task Force, AVIATION FUEL

SAFETY - 1975, CRC Project No. CA-52-74, Coordinating Reearch Council, Inc., New York, November 1975.

4. Marcy, J. F., A STUDY OF AIR TRANSPORT PASSENGER CABIN

FIRES AND MATERIALS, Federal Aviation Agency; Technical

Report FAA-ADS-44, National Aviation Agency, National

Aviation Facilities Experimental Center, Atlantic City,

New Jersey, December 1965, AD 654542.

5. Heins, D., and Brenneman, J., THE FIRE TEST RESULTS, The

Airline Pilot, Vol. 35, No. 10, October 1966, pp. 8-lT,-

6. Mohler, S. R., AIR CRASH SURVIVAL8 INJURIES ?ND EVACUA- TION TOXIC HAZARDS, Aviation, Space, and Environmental

Medicine, January 1975T, pp. 86-88.

7. Pesman, G. J., APPRAISAL OF HAZARDS TO HUMAN SURVIVAL IN

AIRPLANE CRASH FIRES, NACA Technical Note 2996, Lewis

Flight Propulsion Laboratory, National Advisory Committee

for Aeronautics, Cleveland, Ohio, September 1953.

8. Buettner, K., Ph.D., EFFECTS OF EXTREME HEAT ON MAN, Journal of the American Medical Association, Vol. 144, OcEo•e

1950, pp. 732-740.

9. Moritz, A. R., M.D., et al., AN EXPLORATION OF THE CASUALTY PRODUCING ATTRIBUTES OF CONFLAGRATIONS: LOCAL AND

SYSTEMIC EFFECTS OF GENERAL CUTANEOUS EXPOSURE TO EXCESSIVE HEAT OF VARYING DURATION AND INTENSITY, Archives of

Pathology, Vol. 43, 1947, pp. 466-502.

10. Pryer, A. J., and Yuill, C. H., MASS FIRE LIFE HAZARD,

Southwest Research Institute, San Antonio, Texas, September 1966.

200

*. . . ' ' ' ', , , ,i i T
REFERENCES (Continued)

11. Forbes, W. H., Sargent, F., and Roughton, F. J. W., THE

RATE OF CARBON MONOXIDE UPTAKE BY NORMAL MEN, American

Journal of Physiology,.Vol. 143, April 1945.

12. FIRE SAFETY ASPECTS OF POLYMERIC MATERIALS, VOLUME 6 -

AIRCRAFT: CIVIL AND MILITARY, Publication NMAB 3186,

National Materials Advisory Board, National Academy of

Sciences, Washington, D. C., 1977, p. 184.

13. Higgins, E. A., Ph.D., et al., THE ACUTE TOXICITY OF BRIEF

EXPOSURES TO HF, HCl, NO AND HCN SINGLY AND IN COMBINATION WITH CO, FAA Civil Aeromedical Institute; Report No.

FAA-AM-71-41, Department of Transportation, Federal Aviation Administration, Office of Aviation Medicine, Washington, D. C., November 1971, AD 735160.

14. Robertson, S. H., and Turnbow, J. W., A METHOD FOR SELECTING A CRASHWORTHY FUEL SYSTEM DESIGN, paper presented at

NATO/A~aRD Operational Helicopter Aviation Medicine Symposium at Fort Rucker, Alabama in May 1978, North Atlantic

Treaty Organization, Advisory Group for Aerospace and Development, Neuilly-sur-Seine, France.

15. Military Specification, MIL-T-27422B, TANK, FUEL, CRASHRESISTANT, AIRCRAFT, Department of Defense, Washington,

D. C., 24 February 1970.

16. Robertson, S. H., and Turnbow, J. W., AIRCRAFT FUEL TANK

DESIGN CRITERIA, Aviation Safety Engineering and Research

of Flight Safety Foundation; USAAVLABS Technical Report

66-24, U. S. Army Aviation Materiel Laboratories, Fort

Eustis, Virginia, March 1966, AD 631610.

17. Cook, R. L., et al., IMPROVED CRASH-RESISTANT FUEL CELL

MATERIAL, Goodyear Aerospace; USAAVLABS Technical Report

67-6, U. S. Army Aviation Materiel Laboratories, Fort

Eustis, Virginia, April 1967, AD 813165L.

18. Robertson, S. H., DEVELOPMENT OF A CRASH-RESISTANT FLAMMABLE FLUIDS SYSTEM FOR THE UH-lA HELICOPTER, Dynamic

Science; USAAVLABS Technical Report 68-82, U. S. Army

Aviation Materiel Laboratories, Fort Eustis, Virginia,

January 1969, AD 688165.

19. Militar Specification, MIL-H-83282A, HYDRAULIC FLUID,

FIRE RESISTANT SYNTHETIC HYDROCARBON BASE, AIRCRAFT, Department of Defense, Washington, D. C., 22 February 1974.

201

6,
REFERENCES (Continued)

20. Weatherford, W. D. Jr., and Wright, B. R., STATUS OF RESEARCH OF ANTIMIST AIRCRAFT TURBINE ENGINE FUELS IN THE

UNITED STATES, in Aircraft Fire Safety, AGARD Conference

Proceedings No. 166, North Atlantic Treaty Organization,

Advisory Group for Aerospace Research and Development,

Neuilly-sur-Seine, France, October 1975, pp. 2.1-2.12,

AD A018180.

21. San Miguel, A., ANTIMISTING FUEL KINEMATICS RELATED TO

AIRCRAFT CRASH LANDINGS, Journal of Aircraft, Vol. 15,

No. 3, March 1978, pp. 137-142.

22. Shaw, L. M., SAFETY EVALUATION OF EMULSIFIED FUELS, Dynamic Science; USAAMRDL Technical Report 71-29, Eustis

Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, June 1971,

AD 729330.

23. Shaw, L. M., SAFETY EVALUATION OF ANTIMIST FUELS, Dynamic

Science; Report 9130-73-112, U. S. Army Mobility Equipment

Research and Development Center, Aberdeen Proving Ground,

Maryland, November 1973, AD 773035.

24. Robertson, S. H., et al., THEORY, DEVELOPMENT, AND TEST

OF A CRASH-FIRE-INERTING SYSTEM FOR RECIPROCATING ENGINE

/ HELICOPTERS, Aviation Safety Engineering and Research of

Flight Safety Foundation; TRECOM Technical Report 63-49,

U. S. Army Transportation Research Command, Fort Eustis,

Virginia, December 1963.

25. Pinkel, I. I., et al., ORIGIN AND PREVENTION OF CRASH

FIRES IN TURBOJET AIRCRAFT, NACA Report 1019, Lewis Flight

Propulsion Laboratory, National Advisory Committee for

Aeronautics, Cleveland, Ohio, 1958.

26. Black, D. 0., CRASH-FIRE PROTECTION SYSTEM FOR A J57 TURBOJET ENGINE USING WATER AS A COOLING AND INERTING AGENT,

NASA Technical Note D-274, National Aeronautics and Space

Administration, Washington, D. C., February 1969.

27. Pinkel, I. I., et al., MECHANISM OF START AND DEVELOPMENT

OF AIRCRAFT CRASH FIRES, NACA Technical Note 2996, Lewis

Flight Propulsion Laboratory, National Advisory Committee

for Aeronautics, Cleveland, Ohio, 1953.

202
S28. Klueg, E. P., et al., AN INVESTIGATION OF IN-FLIGHT FIRE

PROTECTION WITH A TURBOFAN POWERPLANT INSTALLATION, Report

No. NA-69-26, Department of Transportation, Federal Aviation Administration, National Aviation Facilities Experimental Center, Atlantic City, New Jersey, April 1969, A

AD 686045.

29. Jones, R. B., et al., AN ENGINEERING STUDY OF AIRCRAFT

CRASH FIRE PREVENTION, Walter Kidde and Company, Inc.;

Technical Report 57370, Wright Aeronautical Development

Center, Ohio, June 1958, AD 155846.

30. Campbell, J. A., APPRAISAL OF THE HAZARDS OF FRICTION

SPARK IGNITION OF AIRCRAFT CRASH FIRES, NACA Technical

Note 4024, Lewis Flight Propulsion Laboratory, National

Advisory Committee for Aeronautics, Cleveland, Ohio, May

1957.

31. Moser, J. D., and Black D. 0., PROPOSED INITIATING SYSTEM FOR CRASH-FIRE PREVENTION SYSTEMS, NACA Technical Note

3774, National Advisory Committee for Aeronautics, Cleveland, Ohio, December 1956.

32. Drummond, J. K., STUDY TO DETERMINE THE APPLICATION OF

AIRCRAFT IGNITION-SOURCE CONTROL SYSTEMS TO FUTURE ARMY

AIRCRAFT, Dynamic Science; USAAMRDL Technical Report

71-35, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia,

June 1971, AD 729870.

33. Kuchta, J. M., FIRE AND EXPLOSION MANUAL FOR AIRCRAFT ACCIDENT INVESTIGATORS, Pittsburg Mining and Safety Research

Center, Bureau of Mines; AFAPL Technical Report 73-74,

Air Force Aero Propulsion Laboratory, Wright-Patterson

Air Force Base, Ohio, August 1973, AD 771191.

34. Smith, E. E., PRODUCT FIRE HAZARD EVALUATION, Journal of

Fire & Flainmability/Consumer Product Flammability, Vol. 2,

"March 1975, pp. 58-69.

35. Brauman, S. K., EFFECT OF SAMPLE TEMPERATURE ON COMBUSTION

PERFORMANCE OF POLYMERS, Journal of Fire & Flammability,

Vol. 8, April 1977, pp. 210-224.

203

C ,777'

..... ...
REFERENCES (Continued)

36. Hilado, C. J., and Cumming, H. J., HYDROCARBON CONCENTRATIONS IN FIRE TOXICITY TESTS AS AN INDICATION OF FLASH

FIRE PROPENSITY, Journal of Fire & Flammability, Vol. 8,

April 1977, pp. 235-240.

37. Gross, D., et al., SMOKE AND GASES PRODUCED BY BURNING

AIRCRAFT INTERIOR MATERIALS, Building Science Series 18,

U. S. Department of Commerce, National Bureau of Standards,

Washington, D. C., February 1969.

38. Einhorn, I. N., PHYSIO-CHEMICAL STUDY OF SMOKE EMISSION

BY AIRCRAFT INTERIOR MATERIALS, PART I: PHYSIOLOGICAL

AND TOXICOLOGICAL ASPECTS OF SMOKE DURING FIRE EXPOSURE,

University of Utah; Report No. FAA-RD-73-50-1, Department of Transportation, Federal Aviatio" Administration,

National Aviation Facilities Experiment Center, Atlantic City, New Jersey, July 1973, AD 7636 2.

39. Spurgeon, J. C., Speitel, L. C., and Feher, R. E., OXIDATIVE PYROLYSIS OF AIRCRAFT INTERIOR MATERIALS, Journal of

Fire & Flammability, Vol. 8, July 1977, pp. 349-363.

40. Spurgeon, J. C., A PRELIMINARY COMPARISON OF LABORATORY

METHODS FOR ASSIGNING A RELATIVE TOXICITY RANKING TO AIRCRAFT INTERIOR MATERIALS, Report No. FAA-RD-75-37, Department of Transportation, Federal Aviation Administration,

National Aviation Facilities Experimental Center, Atlantic

City, New Jersey, October 1975, AD A018148.

41. Hilado, C. J., FLAMMABILITY TESTS, 1975: A REVIEW, Fire

Technology, Vol. 11, No. 4, November 1975, pp. 282-T9-.

42. Federal Aviation Regulations, AIRWORTHINESS STANDARDS:

TRANSPORT CATEGORY AIRPLANES, Part 25, Section 25.853.

43. PROPOSED GUIDELINES FOR FLAMMABILITY AND SMOKE EMISSIONS

SPECIFICATIONS (Unofficial), Department of Transportation,

Transportation Systems Center, Cambridge, Massachusetts.

44. FIRE SAFETY ASPECTS OF POLYMERIC MATERIALS, VOLUME 6 -

AIRCRAFT: CIVIL AND MILITARY, National Materials Advisory

Board, National Academy of Sciences, Washington, D. C.,

1977, p. 11.

45. Fish, R. H., AMES T-3 FIRE TEST FACILITY - AIRCRAFT CRASH

FIRE SIMULATION, Journal of Fire & Flammability, Vol. 7,

October 1976, pp. 470-481.

204

........... u
4' REFERENCES (Continued)

46. Federal Aviation Administration, Proposed Regulations,

TRANSPORT CATEGORY AIRPLANES: SMOKE EMISSION FROM COMPARTMENT INTERIOR MATERIALS, Federal Register, 40

F.R.6506, 12 February 1975.

47. Crane, C. R., Ph.D., et al., INHALATION TOXICOLOGY: I.

DESIGN OF A SMALL-ANIMAL TEST SYSTEM: II. DETERMINATION

OF THE RELATIVE TOXIC HAZARDS OF 75 AIRCRAFT CABIN MATERIALS, FAA Civil Aeromedical Institute; Report No. 'AA-AM77-9, Department of Transportation, Federal Aviation Administration, Office of Aviation Medicine, Washington,

D. C., March 1977, AD A043646.

48. Hilado, C. J., FIRE RESPONSE CHARACTERISTICS OF NONMETALLIC MATERIALS: A REVIEW OF RECENT PAPERS AND REPORTS,

Journal of Fire & Flammability, Vol. 7, October 1976, pp. 539-558...

49. Nelson# G. L., et al., MATERIAL PERFORMANCE IN TRANSPORTATION VEHICLE INTERIORS, Journal of Fire & Flammability,

Vol. 8, July 1977, pp. 262-278.

50. Dusken, F. E., FIRE TESTING OF AIRCRAFT CABINS, Journal

of Fire & Flammability, Vol. 8, April 1977, pp. 193-201.

51. Marcy, J. F., and Johnson, R., FLAMING AND SELF-EXTINGUISHING CHARACTERISTICS OF AIRCRAFT CABIN INTERIOR MATERIALS, Report No. NA-68-30 [DS-68-13], Department of

Transportation, Federal Aviation Administration, National

Aviation Facilities Experimental Center, Atlantic City,

New Jersey, July 1968, AD 673084.

52. Einhorn, I. M., Kanakia, M. D., and Seader, J. D., PHYSIOCHEMICAL STUDY OF SMOKE EMISSION BY AIRCRAFT INTERIOR MATERIALS, Part II: RIGID- AND FLEXIBLE-URETHANE FOAMS,

University of Utah; Report No. FAA-RD-73-50,II, Department

of Transportation, Federal Aviation Administration, National Aviation Facilities Experimental Center, Atlantic

City, New Jersey, July 1973, AD 763935.

" 53. Fewell, L. L., et al., eds., CONFERENCE ON THE DEVELOPMENT

OF FIRE-RESISTANT AIRCRAFT PASSENGER SEATS, NASA Technical

Memorandum X-73144, National Aeronautics and Space Administration, Ames Research Center, Moffett Field, California,

August 1976, pp. 115-125.

54. Morford, R. H., THE FLAMMABILITY OF NEOPRENE CUSHIONING

FOAM, Journal of Fire & Flammability, Vol. 8, July 1977,

pp. 279-299.

(. 205

AI
REFERENCES (Continued)

55. Hilado, C. J., THE CORRELATION OF LABORATORY TEST RESULTS

WITH BEHAVIOR IN REAL FIRES, Journal of Fire & Flammability, Vol. 8, April 1977, pp. 202-209.

56. Parker, J. A., Ph.D., POLYPHOSPHAZENE SEAT CUSHION APPLICATIONS, American Research Center; in Conference on the

Development of Fire-Resistant Aircraft Passenger Seats,

Fewell, L. L., et al., eds., NASA Technical Memorandum

X-73144, National Aeronautics and Space Administration,

Washington, D. C., August 1976.

57. Gagliani, J., FIRE RESISTANT RESILIENT FOAMS - FINAL REPORT, Solar Division, International Harvester; Report No.

N76-18278, National Aeronautics and Space Administration,

Lyndon B. Johnson Space Center, Houston, Texas, February 1976.

58. Tesoro, G. C., Ph.D., FABRICS FOR FIRE RESISTANT PASSENGER

SEATS IN AIRCRAFT, Massachusetts Institute of Technology;

in Conference on Fire Resistant Materials (Firemen),

Kourtides, D. A., ed., NASA Technical Memorandum 78523, National Aeronautics and Space Administration, Ames Research

Center, Moffett Field, California, October 1978.

59. Tesoro, G. C., Ph.D., STATE OF THE ART MATERIALS USED IN

AIRCRAFT PASSENGER SEATS, Massachusetts Institute of Technology; in Conference on the Development of Fire-Resistant

Aircraft Passenger Seats, Fewell, L. L., et al., eds., NASA

Technical Memorandum X-73144, National Aeronautics and

Space Administration, Washington, D. C., August 1976.

60. Silverman, B., AIRCRAFT INTERIOR THERMOPLASTIC MATERIALS,

Lockheed Missiles and Space Company; in Conference on the

Development of Fire-Resistant Aircraft Passenger Seats,

Fewell, L. L., et al., eds., NASA Technical Memorandum

X-73144, National Aeronautics and Space Administration,

Washington, D. C., August 1976.

61. Parker, J. A., Ph.D., et al., FIRE DYNAMICS OF MODERN

AIRCRAFT FROM A MATERIALS POINT OF VIEW, in Aircraft Fire

Safety, AGARD Conference ProcMeedings No. 166, North Atlantic Treaty Organization, Advisor.cy Grou-pfor Aerospace Research and Developmont, Neuilly-sur-Seine, France, October 1975, pp. 10.1-10.11, AD A018180.

"206
j

REFERENCES (Continued)

62. Anderson, R. A., et al., EVALUATION OF MATERIALS AND CONCEPTS FOR AIRCRAFT FIRE PROTECTION, Boeing Commercial Airplane Company; Report No. NASA CR-137838, National Aero- nautics and Space Administration, Ames Research Center, Moffett Field, California, April 1976.

63. Atallah, S., and Buccigross, H., INVESTIGATION AND EVALUATION OF NONFLAMMABLE, FIRE-RETARDANT MATERIALS, Arthur

D. Little, Inc.; USAAMRDL Technical Report 72-52, Eustis

Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, November 1972,

AD 906699L.

64. Saczal3ki, K., et al., eds. AIRCRAFT CPAS-WORTHINESS, Charlottesville, Virginia, University Press of Virginia, 1975, pp. 121-139.

65. Flaig, J. W., HELICOPTER FLOTATION AND PERSONNEL SAFETY

IN UNPLANNED WATER LANDING FROM FY 1960 THROUGH FY 1974, •I Letter Report 75-61, Naval Air Systems Command# Washing- ton, D. C., June 1975.

66. Rice, E. V., and Greear, J. F. III, UNDERWATER ESCAPE FROM

HELICOPTERS, Naval Safety Center, Norfolk, Virginia; paper

presented at the l1th Annual Symposium of the Survival

and Flight Equipment Association, Phoenix, Arizona, Octo- ber 7-11, 1973.

67. Ryack, B. L., et al., THE EFFECTIVENESS OF ESCAPE HATCH

ILLUMINATION AS AN AID TO EGRESS FROM A SUBMERGED HELICOPTER, Report No. 856, Naval Submarine Medical Research Laboratory, Groton, Connecticut, November 1977.

68. Smith, P. F., Lurial, S. M., Ryack, B. L., LUMINANCETHRESHOLDS OF THE WATER-IMMERSED EYE, Report No. 857, Naval Submarine Medical Research Laboratory, Groton, Connecticut, March 1978.

69. Saczalski, K., at al., ads., AIRCRAFT CRASHWORTHINESS,

Charlottesville, Virginia, University Press of Virginia,

1975, pp. 645-667.

70. Veghte, J. D., COLD SEA SURVIVAL, Aerospace Medicine, 4

Vol. 43, No. 5, May 1972.

71. DESIGN NOTE 3Q6, in AFSC: DHL-6, System Safety, 4th Edition, Air Force Systems Command, Wright-Patterson Air

Force Base, Ohio, July 1974.

207
REFERENCES (Continued)

72. McFadden and Swearingen, FORCES THAT MAY BE EXERTED BY MAN

IN THE OPERATION OF AIRCRAFT DOOR HANDLES, Journal of the

Human Factors Society of America, Vol. 1, No. 1, September

1958.

73. Nicholson, D. E., and Burkdoll, F. B., AN EMERGENCY LIFESAVING INSTANT EXIT SYSTEM FOR CARGO, CARGO-TRANSPORT AND

PASSENGER AIRCRAFT, Explosive Technology; ASD Technical

Report 71-41, Aeronautical Systems Division, Air Force

Systems Command, Wright-Patterson Air Force Base, Ohio,

June 1971, AD 736056.

74. Federal Aviation Regulations, AIRWORTHINESS STANDARDS:

TRANSPORT CATEGORY ROTORCRAFT, Part 29, Section 29.811.

75. Military Specification, MIL-L-6503H, LIGHTING EQUIPMENT,

AIRCRAFT, GENERAL SPECIFICATION rOR INSTALLATION OF, Department of Defense, Washington, D. C., 2 April 1975.

76. Mendenhall, C. E., et al., COLLEGE PHYSICS, 3rd Edition,

Boston, Massachusetts, D. C. Heath and Company, 1950,

p. 476.

77. Garner, J. D., and Lowrey, D. L., EXIT SIGN COMPARISONS

IN CLEAR AIR AND SMOKE, unpublished paper presented at an A-20 meeting, Society of Automotive Engineers, New York,

December 1976.

78. Federal Aviation Regulations, AIRWORTHINESS STANDARDS:

TRANSPORT CATEGORY AIRPLANES, Part 25, Section 25.811.

79. Technical Bulletin 746-93-2, PAINTING AND MARKING OF ARMY

AIRCRAFT, Department of the Army, Washington, D. C., January 1971.

80. EMERGENCY LOCATOR TRANSMITTERS: AN OVERVIEW, Report No.

NTSB-AAS-78-1, National Transportation Safety Board, Washington, D. C., 26 January 1978. S81. Johnson, N., and Sanderson, S., SPILLED FUEL IGNITION

SOURCES AND COUNTERMEASURES, Ultrasystems, Inc., Dynamic

Science; Report No. DOT-HS-801 722, U. S. Department of

Transportation, National Highway Traffic Safety Administration, Washington, D. C., September 1975. V 82. Federal Aviation Technical Order, EMERGENCY LOCATOR TRANSMITTERS, (TSO) C91.

208

'• ,• 2 o
REFERENCES (Continued)

83. MINIMUM PERFORMANCE STANDARDS - EMERGENCY LOCATOR TRANSMITTERS, Document No. DO-147, Radio Technical. Commission

for Aeronautics, Washington, D. C., 5 November 1970.

84. DEVELOPMENT OF ELT CRASH SENSOR PERFORMANCE SPECIFICATION

AND TEST PROCEDURES - FINAL REPORT, Crash Research Institute, Robertson Research, Inc.; Contract No. DOT-FA76WA3842, Department of Transportation, Federal Aviation Administration, Washington, D. C., June 1977.

85. Military Standard, MIL-STD-810B, ENVIRONMENTAL TEST METHODS,

Department of Defense, Washington, D. C., 11 June 1967.

209

iI )

I~i I
BIBLIOGRAPHY

Aerospace Recommended Practice, AIRCRAFT EVACUATION SIGNAL SYSTEM, ARP 1178, Society of Automotive Engineers, Inc., New York,

July 1971.

Aerospace Recommended Practice, COMBINATION EVACUATION SLIDE/

LIFE RAFT, ARP 1282, Society of Automotive Engineers, Inc.,

New York, July 1970.

Aerospace Recommended Practice, EMERGENCY EVACUATION ILLUMINATION, ARP 503B, Society of Automotive Engineers, Inc.,, New York,

Revised August 1970.

Aerospace Recommended Practice, EMERGENCY PLACARDING - INTERNAL

AND EXTERNAL, ARP 577A, Society of Automotive Enginee:rs, Inc.,

New York, Revised October 1962.

Aerospace Recommended Practice, ESCAPE PROVISIONS, FLIGHT DECK,

ARP 808A, Society of Automotive Engineers, Inc., New York, Revised December 1970.

Aerospace Recommended Practice, EXITS AND THEIR OPERATION - AIR

TRANSPORT CABIN EMERGENCY, ARP 488B, Society of Automotive Engineers, Inc., New York, Revised October 1973.

Aerospace Recommended Practice, PASSENGER CABIN EMERGENCY EQUIPMENT STOWAGE, ARP 997, Society of Automotive Engineers, Inc.,

New York, August 1967.

Aerospace Recommended Practice, STOWAGE OF FLIGHT CREW'S SURVIVAL EMERGENCY AND MISCELLANEOUS EQUIPMENT, ARP 917, Society

of Automotive Engineers, Inc., New York, March 1966.

Aerospace Recommended Practice, SURVIVAL KIT -- LIFE RAFTS AND

SLIDE RAFTS, ARP 1282, Society of Automotive Engineers, Inc.,

New York, July 1973.

Andersoii, M. C., and Burkdoll, F. B., AN EMERGENCY LIFE-SAVING,

INSTANT EXIT SYSTEM FOR CARGO, CARGO-TRANSPORT AND IPASSENGER

AIRCRAFT, VOLUME II, Explosive Technology; ASD Technical Report

71-41, Aeronautical Systems Division, Air Force Systems Command,

Wright-Patterson Air Force Base, Ohio, February 1974, AD 778824.

Anson, B., EVALUATION OF SELF-SEALING BREAKAWAY VALVES FOR

CRASHWORTHY AIRCRAFT FUEL SYSTEMS, Dynamic Science Engineering

Operations; USAAMRDL Technical Report 71-65, Eustis Directorate,

U. S. Army Air Mobility Research and Development Laboratory,

Fort Eustie, Virginia, November 1971, AD 738204.

210

/ •, I,• 41•

. I . i I a I l I I I
BIBLIOGRAPHY (Continued)

Arnold, D. B., Burnside, J. V., and Hajari, J. V., DEVELOPMENT

OF LIGHTWEIGHT FIRE-RETARDANT, LOW-SMOKE, HIGH-STRENGTH, THERMALLY STABLE AIRCRAFT FLOOR PANELING, Boeing Commercial Airplane Company; Technical Report NASA-CR-147750, National Aeronautics and Space Administration, Houston, Texas, April 1976.

Asher, W. J., CONCEPTUAL DESIGN STUDY OF DEMULSIFICATION OF

FUEL EMULSIONS BY A COALESCENCE PROCESS, Esso Research and Engineering Co.; USAAVLABS Technical Report 69-16, U. S. Army

Aviation Materiel Laboratories, Fort Eustis, Virginia, April

1969, AD 688973.

Atkinson, A. J., EVALUATION OF EXPERIMENTAL SAFETY FUELS IN A

CONVENTIONAL GAS TURBINE COMBUSTION SYSTEM, Naval Air Propulsion Test Center; Report No. FAA-NA-69-1 (DS-68-27), Department

of Transportation, Federal Aviation Administration, Washington,

D. C., April 1969, AD 686840.

Bruggink, G., EMERGENCY LANDING AND DITCHING TECHNIQUES IN HELICOPTERS, U. S. Army Board for Aviation Accident Research, Fort

Rucker, Alabama, January 1968.

Byerley, D. L., EVALUATION OF VARIOUS SIZES AND CONFIGURATIONS

OF FUEL TANKS, Firestone Tire & Rubber Co.; USAAMRDL Technical Report 73-74, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratories, Fort Eustis, Virginia,

December 1973, AD 777801.

Carhart, H. W., Pinkel, I. I., Warren, J. H., AIRCRAFT KEROSENE

VERSUS WIDE-CUT FUEL-SAFETY CONSIDERATIONS, paper presented at

Air Transportation Meeting at New York City, Society of Automotive Engineers, New York, May 1976.

Catiller, J. B., Major, USAF, ed., AIRCRAFT FIRE SAFETY, AGARD

Conference Proceedings No. 166, North Altantic Treaty Organization, Advisory Group for Aerospace Research and Development,

Neuilly-sur-Seine, France, October 1975, AD A018180.

Cook, R. L., et al., EVALUATION OF THE UH-ID/H HELICOPTER

CRASHWORTHY FUEL SYSTEM IN A CRASH ENVIRONMENT, Dynamic Science Engineering Operations, Phoenix, Arizona, November 1971.

Federal Aviation Administration, AIR CARRIER CABIN SAFETY

A SURVEY, Department of Transportation, Washington, D. C., December 1976, AD A037906. •

Field Manual 20-151, AIRCRAFT EMERGENCY PROCEDURES OVER WATER,

Department of the Army, Washington, D. C., November 1968.

'" 211

., . .... ...... 
BIBLIOGRAPHY Continued)

FIRE SAFETY ASPECTS OF POLYMERIC MATERIALS, VOLUME 6 - AIRCRAFT:

CIVIL AND MILITARY, Publication NMAB 318-6, National Materials

Advisory Board, National Academy of Sciences, Washington, D. C.,

1977.

Geyer, G. B., EFFECT OF GROUND CRASH FIRE ON AIRCRAFT FUSELAGE

INTEGRITY, Report No. NA-69-37, Department of Transportation,

Federal Aviation Administration, National Aviation Facilities

Experimental Center, Atlantic City, New Jersey, December 1969,

AD 698806.

Gilewicz, A. E., Major, USA, ASSESSMENT OF THE BENEFITS OF AIRCRAFT CRASHWORTHINESS, USAAAVS System Safety Newsletter, U. S.

Army Agency for Aviation Safety, Vol. 5, No. 3, 1976.

Hall, J. F., PREDICTION OF TOLERANCE IN COLD WATER AND LIFE

RAFT EXPOSURES, Aerospace Medicine, Vol. 43, No. 3, March 1972.

Harris, J. C., and Steinmetz, E. A., OPTIMIZATION OF JP-4 FUEL

EMULSIONS AND DEVELOPMENT OF DESIGN CONCEPTS FOR THEIR DEMULSIFICATION, Monsanto Research Corporation; USAAVLABS Technical

Report 68-79, U. S. Army Aviation Materiel Laboratories, Fort

Eustis, Virginia, November 1968, AD 683597.

Harvey, E. A., and Jensen, D. L., PRECISION SURVIVOR LOCATOR

FOR SEARCH AND RESCUE APPLICATIONS, paper presented at 29th

Annual National Forum of the American Helicopter Society in

Washington, D. C., American Helicopter Society, New York, May

1973.

Harvey, T., and Monarch, J., ENVIRONMENTAL TESTING OF A GAS TURBINE ENGINE UTILIZING EF4-101 EMULSIFIED JP-4 FUEL, Continental

Aviation and Research Corporation; USAAVLABS Technical Report

68-55, U. S. Army Aviation Materiel Laboratories, Fort Eustis,

Virginia, August 1968, AD 674205.

Hibbard, R. P, and Hacker, P. T., AN EVALUATION OF THE RELATIVE FIRE HAZARDS OF JET A AND JET B FOR COMMERCIAL FLIGHT,

NASA Technical Memorandum X-71437, Lewis Research Center, National Aeronautics and Space Administration, Washington, D. C.,

October 1973.

Hilado, C. J., SMOKE FROM CELLULAR POLYMERS, Fire Technology,

May 1969, p. 130.

Institute for Research, Inc., DEVELOPMENT OF A PROTECTIVE,

DECORATIVE, FIRE-RESISTANT, LOW-SMOKE EMITTING, THERMALLY

STABLE COATING MATERIAL, Report No. N 76-18234, Houston, Texas,

January 1976.

212
/, BIBLIOGRAPHY (Continued)

Johnson, N. B., CRASHWORTHY FUEL SYSTEM DESIGN CRITERIA AND

ANALYSIS, Dynamic Science; USAAVLABS Technical Report 71-8,

Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratories, Fort Eustis, Virginia, March 1971,

AD 723988.

Kosvic, T. C., Holgeson, N. L., and Breen, B. P., FLIGHT VIBRATION AND ENVIRONMENTAL EFFECTS ON FORMATION OF COMBUSTIBLE MIXTURES WITHIN AIRCRAFT FUEL TANKS, Dynamic Science; USAAVLABS

Technical Report 70-43, U. S. Army Aviation Materiel Laboratories, Fort Eustis, Virginia, September 1970, AD 875901.

Kourtides, D. A., and Parker, J. A., inventors, LOW DENSITY

BISMALEIMIDE-CARBON MICROBALLOON COMPOSITES, NASA Case No.

ARC-11040-1, National Aeronautics and Space Administration,

Ames Research Center, Moffett Field, California, 16 March

1977.

Malmberg, Q. C., and Wiggins, E. W., ADVANCED FLAME ARRESTOR

MATERIALS AND TECHNIQUES FOR FUEL TANK PROTECTION, McDonnell

Douglas Corp.; Technical Report 72-12, Air Force Aero Propulsion Laboratory, Wright-Patterson Air Force Base, Ohio, March

1972, AD 743016.

Malmberg, Q. C., INTEGRATED AIRCRAFT FUEL TANK FIRE AND EXPLOSION PROTECTION SYSTEMS-PHASE I AND II, McDonnell Douglas Corp.,

St. Louis, Missouri, AD A020589.

Marcy, J. F., et al., FLAMMABILITY AND SMOKE CHARACTERISTICS

OF AIRCRAFT INTERIOR MATERIALS, Technical Report FAA-ADS-3,

Department of Transportation, Federal Aviation Agency, National Aviation Facilities Experimental Center, Atlantic City, New

Jersey, January 1964, AD 600387.

•I Mason, R. V., SMOKE AND TOXICITY HAZARDS IN AIRCRAFT CABIN

FURNISHINGS, Annual of Occupational Hygiene, Vol. 17, 1974, l* • pp. 159-165.,,

McDonald, W. C., NEW ENERGY-ABSORBING MATERIALS FOR CRASHRESISTANT FUEL TANKS, Goodyear Aerospace Corporation; presented at National Business Aircraft Meeting, Wichita, Kansas,

Society of Automotive Engineers, New York, April 1968.

McFadden, E. B., and Simpson, J. M., FLOTATION CHARACTERISTICS

OF AIRCRAFT-PASSENGER SEAT CUSHIONS, Federal Aviation Agency,

Washington, D. C., April 1966.

213

Ali
BIBLIOGRAPHY (Continued)

Miller, R. E., and Wilford, S. P., THE DEVELOPMENT OF AN AIRCRAFT SAFETY FUEL, Royal Aircraft Establishment, England; Report No. A76-14419, American Institute of Aeronautics and Asstronautics, New York, April 1975.

Nissley, P. M., Held, T. L., STRUCTURAL DESIGN FOR FUEL CONTAINMENT UNDER SURVIVABLE CRASH CONDITIONS, General Dynamics/

Convair, FAA Technical Report ADS-19, Department of Transportation, Federal Aviation Administration, Washington, D. C.,

August 1964.

Pedriani, C. M., A STUDY OF THE FUEL/AIR VAPOR CHARACTERISTICS IN THE ULLAGE OF AIRCRAFT FUEL TANKS, USSAVLABS Technical

Note 3, U. S. Army Aviation Materiel Laboratories, Fort Eustis,

Virginia, June 1970, AD 873252.

Pedriani, C. M., DESIGN CRITERIA FOR PREVENTION OF FIRES FROM

API IMPACTS IN VOIDS ADJACENT TO HELICOPTER FUEL TANKS,

USAAMRDL Technical Note 20, Eustis Directorate, U. S. Army Air

Mobility Research and Development Laboratory, Fort Eustis, Virginia, August 1975, AD B006677L.

Pedriani, C. M., EXPERIMENTAL DETERMINATION OF THE IGNITION

LIMITS OF JP-4 FUEL WHEN EXPOSED TO CALIBER .30 INCENDIARY PROJECTILES, USAAMRDL Technical Report 71-48, Eustis Directorate,

U. S. Army Air Mobility Research and Development Laboratory,

Fort Eustis, Virginia, July 1971, AD 730343.

Pedriani, C. M., INCENDIARY GUNFIRE EVALUATION OF A CONDENSATEFORMED FUEL FOG INERTING SYSTEM, USAAMRDL Technical Note 15,

Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, May 1974, AD 780837.

Perrella, W. M., Jr., TESTS OF CRASH-RESISTANT FUEL SYSTEM FOR

GENERAL AVIATION AIRCRAFT, Department of Transportation, Federal Aviation Administration, Washington, D. C., March 1978.

Reed, T. 0., EFFECTS OF POLYURETHANE FOAM ON FUEL SYSTEM CONTAMINATION, Technical Memorandum ENJI-70-10, Fluid Systems

Branch, Directorate of Propulsion and Power Subsystems Engineering, Wright-Patterson Air Force Base, Ohio, November 1970.

Roebuck, J. A., Jr., NEW CONCEPTS FOR EMERGENCY EVACUATION OF

TRANSPORT AIRCRAFT FOLLOWING SURVIVABLE ACCIDENTS, North American Rockwell Corporation; Report No. ADS-68-2, Department of

Transportation, Federal Aviation Administration, Washington,

D. C., January 1968, AD 665329.

214
BIBLIOGRAPHY (Continued)

Ross, J. H., RECENT DEVELOPMENTS IN COMMERCIAL FIRE-RESISTANT

"FIBROUS MATERIALS, Safe Engineering, 4th Quarter-Midwinter, V. 1972-73, pp. 16-21.

Saczalski1 K., et al., eds., AIRCRAFT CRASHWORTHINESS, Charlotteville, Virginia, University Press of Virginia, 1975.

Salmon, R. F., STUDY OF TURBINE ENGINE OPERATION WITH GELLED

FUELS, Report No. FAA-DS-70-6, Department of Transportation, Federal Aviation Administration, Washington, 0. C., May 1970,

AD 711765.

San Miguel, A., and Williams, M. D., ANTIMISTING FUEL SPILLAGE/

AIR SHEAR TESTS AT NAVAL WEAPONS CENTER, Report No. FAA-RD78-50, Department of Transportation, Federal Aviation Administration, Washington, D. C., March 1978, AD A056113.

Scheuerman, He P., CRASH RESISTANT FUEL SYSTEMS DEMONSTRATIONS AND EVALUATION, Report No. FAA-RD-71-27, National Aviation Facilities Experimental Center, Atlantic City, NqW Jersey, August

1971.

Scull, W. D., RELATION BETWEEN INFLAMMABLES AND IGNITION

SOURCES IN AIRCRAFT ENVIRONMENTS, Lewis Flight Propulsion Laboratory; NACA Report 1019, National Advisory Committee for

Aeronautics, Cleveland, Ohio, 1951.

Smith, H. D., CRASH-RESISTANT FUEL TANKS FOR HELICOPTERS AND

GENERAL AVIATION AIRCRAFT, presented at National Business Aircraft Meeting, Wichita, Kansas, Society of Automotive Engineers,

New York, April 1974.

Snow, Clyde C., et al., SURVIVAL IN EMERGENCY ESCAPE FROM

PASSENGER AIRCRAFT, Civil Aeromedical Institute; Report No.

AM 70-16, Department of Transportation, Federal Aviation Administration, Washington, D. C., October 1970, AD 735388.

Snyder, R. G., ADVANCED TECHNIQUES IN CRASH IMPACT PROTECTION

AND EMERGENCY EGRESS FROM AIR TRANSPORT AIRCRAFT, University

of Michigan; AGARDOGRAPH No. 221, North Atlantic Treaty Organ- ization, Advisory Group for Aerospace Research and Development,

Neuilly-sur-Seine, France, June 1976, AD A029375.

Snyder, R. G., CRASHWORTHINESS INVESTIGATION OF GENERAL AVIATION ACCIDENTS, University of Michigan; presented at National

Business Aircraft Meeting, Wichita, Kansas, Society of Automotive Engineers, New York, April 1975.

215

* I I I-ahIIIIllIIiI|
BIBLIOGRAPHY (Continued)

Snyder, R. G., et al., EMERGENCY DITCHING OF GENERAL AVIATION

AIRCRAFT, Proceedings of the 12th Annual Survival and Flight

Equipment Symposium, September 1974.

Snyder, R. G., U. S. EXPERIENCE IN WATER DITCHING OF CIVIL AIRCRAFT, presented at International Congress of Aviation and

Space Medicine, Beirut, Lebanon, October 1974.

Sommers, J., Jr., Clark, J. H., THE DEVELOPMENT OF AIRCRAFT

CRASH-RESISTANT FUEL CELLS, SAFETY VALVES, AND BREAKAWAY ACCESSORIES, Federal Aviation Agency; presented at National Aeronautics Meeting, Society of Automotive Engineers, Inc., New

York, April 1962.

Spencer, L. E., Captain, USA, BREAKAWAY FUEL CELL CONCEPT, Aviation Crash Injury Research of Flight Safety Foundation; TRECOM

Technical Report 62-37, U. S. Army Transportation Research Command, Fort Eustis, Virginia, May 1962.

Spezia, E., U. S. ARMY HELICOPTER ACCIDENTS INVOLVING FIRE,

Report 1-64, U. S. Army Board for Aviation Accident Research,

Fort Rucker, Alabama, 1963.

Supkis, Daniel E., REFURBISHMENT OF NASA AIRCRAFT WITH FIRERETARDANT MATERIALS, Lyndon B. Johnson Space Center; Technical Memorandum X-58165, National Aeronautics and Space Administration, Washington, D. C., October 1975.

U. S. AIR CARRIER ACCIDENTS INVOLVING FIRE, 1965 THROUGH 1974

AND FACTORS AFFECTING THE STATISTICS, Report No. NTSB-AAS-771-1,

National Transportation Safety Board, Washington, D. C., February 1977.

Villemain, H. F., DESIGN AND EVALUATION OF CRASHWORTHY, BALLISTICALLY TOLERANT FUEL TANK MATERIAL, The Goodyear Tire & Rubber

Company; USAAMRDL Technical Report 74-103, Eustis Directorate,

U. S. Army Air Mobility Research and Development Laboratory,

Fort Eustis, Virginia, January 1975, AD B002702.

Zapp, J. A., Jr., THE TOXICOLOGY OF FIRE, Medical Division

Special Report No. 4, Army Chemical Center, Maryland, April

1951.

Zinn, S. U., Jr., CRASH FIRE HAZARD EVALUATION OF JET FUELS,

Federal Aviation Administration; Paper No. 720324, presented

at National Business Aircraft Meeting at Wichita, Kansas, Society of Automotive Engineers, Inc., New York, March 1972.

216

Ii i~ IL
APPENDIX A

RELATION OF PAIN THRESHOLD TIME TO

HEAT SOURCE TEMPERATURE

The pain threshold curves in Figure 7 (Section 3.3.1.1), which

apply to visible or exposed areas of the skin, were generated

* from data in References Al, A2, A3, A4, A5, A6, and A7. The

curves that account for variable radiating surfaces (F) were

determined in the following manner.

The most significant variables that determine the rate of heat

absorption by heat radiation are: (1) temperature of radiative

source, (2) fraction of visible hemisphere at elevated temperature (F), and (3) emissivity of radiative source. Taking these

factors into consideration, the rate of radiative heat absorp- tion is

qR - aeociF*T4 (A-1)

where a - absorptivity of skin surface

e - emissivity of radiative source

a * Stephan-Boltzmann constant

= 4.88 x 108 kcal m- 2 hr" (oK)f 4 4

F - fraction of visible hemisphere occupied by radiating surface

T - temperature of radiative surface, OK

Assuming that skin absorptivity and source emissivity are both

equal to 0.85, Equation (A-i) becomes

T 42

q. 3.50 F [kcal m 2 hr-J (A-2)

or

_-- - 0.73 (#) 1/4 (q) 1/4 K (A-3)

217

': {'k.
Equation (A-3) relates the temperature of the emitting source

to the rate of heat absorption per unit area by the skin. (The

emitter occupies fraction F of the visible hemisphere.)

Equation (A-3) was used to plot the radiative burn curves for

four cases (F = 1.0, 0.5, 0.25, and 0.10) in Figure 7.

REFERENCES - APPENDIX A

A-i Pesman, G. J., APPRAISAL OF HAZARDS TO HUMAN SURVIVAL IN

AIRPLANE CRASH FIRES, NACA Technical Note 2996, Lewis

Flight Propulsion Laboratory, National Advisory Committee

for Aeronautics, Cleveland, Ohio, September 1953.

A-2 Moritz, A. R., M.D., et al., AN EXPLORATION OF THE CASUALTY PRODUCING ATTRIBUTES OF CONFLAGRATIONS: LOCAL AND

SYSTEMIC EFFECTS OF GENERAL CUTANEOUS EXPOSURE TO EXCESSIVE HEAT OF VARYING DURATION AND INTENSITY, Archives of

Pathology, Vol. 43, 1947, pp. 466-502.

A-3 Bieberdorf, F. W., and Yuill, C. H., AN INVESTIGATION OF

THE HAZARDS OF COMBUSTION PRODUCTS IN BUILDING FIRES,

Southwest Research Institute, San Antonio, Texas, 14 October 1963.

A-4 Buettner, K., Ph.D., EFFECTS OF EXTREME HEAT ON MAN,

Journal of the American Medical Association, Vol. 144,

October 1950, pp. 732-740.

A-5 Henriques, F. C., Jr., Ph.D., THE PREDICTABILITY AND THE

SIGNIFICANCE OF THERMALLY INDUCED RATE PROCESSES LEADING

TO IRREVERSIBLE EPIDERMAL INJURY, Archives of Pathology,

Vol. 43, 1947, pp. 489-502.

A-6 Moritz, A. R., M.D., et al., THE EFFECTS OF INHALED HEAT

ON THE AIR PASSAGES AND LUNGS, American Journal of Pathology, Vol. 21, March 1945, pp. 311-331.

A-7 Moritz, A. R., M.D., et al., THE RELATIVE IMPORTANT OF

TIME AND SURFACE TEMPERATURES IN CAUSATION OF CUTANEOUS

BURNS, American Journal of Pathology, Vol. 23, 1947,

pp. 695-720.

218

'.' 4
INDEX 

!

Acceleration IDX

crash 188

vibration 190, 192

Acrylic glazing 120, 137 : ~Adhesives 129 

i

Aircraft skin 27, 29-30

Aisles 174

Ames T-3 test 117 .Antenna 83, 186, 198

ASTM E 162 radiant panel test 116, 138

F Attenuation coefficient 150-153, 155

Autoignition temperature 104-105

Battery 81-83

Boost pump 71

Cabin Fire Simulator 120

Carbon monoxide

animal versus human response to 119

concentrations of, from fires 31-32

generation of, from burning materials 113

human tolerance to 37-41

Carboxyhemoglobin

escape limit due to 31, 37, 40

levels of, in aircraft fires 31-32

physiological effects of 37-39

Cockpit canopies 163, 170

Compartment doors 174-175

Crash parameters 76, 187-188

Crash locator beacons

antenna for 198

electrical wiring of 199

power supply for 198

sensors for 186-197

transmitter for 197-198

types of 185-186

Crash sensors 98-99, 101, 186-197

Crew chief stations 169, 172, 183

Ditching

aircraft flotation systems for 156-158

definition of 145

egress problems in 147-148

emergency equipment for 158-159 :+i.,+emergency lighting for 149-155

escape openings for 148-149

statistics 145-147

Drains, fuel tank 74

Electrical system design 80-83

219 t+II
INDEX (Continued)

Emergency exits

access to 174-175

explosively cut 149, 170-174

lights for 176-179

location of 164-165

marking of 181-183

number of 162-163

operation of 165-170

release handles for 167-169

size of 161-162

types of 161

Engine ignition sources from 83-87

inerting systems for 87-92 Escape limit 33

Escape time

allowable, for emergency egress 160

from carbon monoxide 31-32

from heat 27-28

Exits. See Emergency exits.

FAA

integrated standard for flammability, smoke, toxic gases 136

interior material tests (FAR 25.853) 116, 136-137

proposed smoke emission requirements 118, 136

requirements for emergency lighting 177

requirements for emergency locator transmitters 198

Fabrics

flammability properties of 127-128

flammability requirements for 137-138

smoke emission requirements for 138-139

Fillur caps 75

Filler necks 71

Filters, fuel 74

Fire endurance 117

Flame retardants 111-114, 120-121, 127

Flame spread rate 106-107, 116

Flammability properties of materials. See Materials, interior.

Flash fire potential 109, 136

Flashover 120

Flight loads 188, 190-193

Flotation

characteristics of aircraft 145-147

systems for aircraft 156-158

Flow diverters 92-95

Foams

flammability requirements for 137-138

for insulation 129-134

for seat cushions 121-127, 135-136

220 

(

L l "A;
INDEX (Continued)

for structural panels 129

smoke emission requirements for 138-139

V" Frangible attachments

design criteria for 53-55

for fuel lines 56

for fuel system components 71-73

for fuel tanks 53-55

for lights 83-84

Fuel

ignition of 27, 80-81, 85-87

"r modification of 78-79

volatility of 27, 79

Fuel lines

construction of 56-64

routing of 64

Fuel system design, general 42-43, 76

Fuel tank

access cover 75

attachments 53-56

design criteria for 49-50

drain 74-75

filler 71-73

fittings 50-53

location 43-44

material 46-50

shape 44-46

specification 46, 47, 50

tests 47-50 Generator 81, 83 

; Glazing 120, 132, 137-138

Heater 96

Heat release rate 108-109, 117

Honeycomb panels 129 

J

Human tolerance

to heat 33-36

to miscellaneous fire factors 41-42

to toxic gases 37-41

Hydraulic fluid 77-79

Hydraulic system design 76-77

Hydrogen cyanide

generation of, in fires 33

generation of, fx.:m burning materials 113, 122

human tolerance to 40-41

toxicity in small animal tests 114, 119

Igniters 82

Ignition

ease of 104-106, 115

sources 80, 83, 87, 96

221

...................---.
INDEX (Continued)

temperature of fuel 85-86

Illumination, emergency 175, 176, 179

Inertia sensors. See Crash sensors.

Inerting agents 88

Inerting systems for electrical systems 80-82 for engines 87-92

for heaters 96

initiating systems for 98-101

Insulation

effect of, on burn-through time 30

"flammability requirements for .135, 137-138

in interior walls 129, 132-134

smoke requirements 138

Intumescent paint 95, 135

Inverter 81, 82

Kynol 127, 128

Life rafts 158-159

Lighting, emergency actuation of 180-181

exterior 179

for ditching 149-155

for emergency exits 176-179

interior 175-176

power source for 180

structure of 179-180

Lights

beacon 83-84

emergency exit 177-179

landing 80

Limiting oxygen index (LOI) 115-116 

:

Luminance 150-155

Magneto 81, 82

IMaterials, interior

flammability of, factors in 104

flammability properties of 139, 1.40-144

flammability tests of 115-117, 119-120, 136-139

' selection criteria for 135-139

MIL-STD-810B 192

MIL-T-27422B 41, 47, 50, 53

NBS smoke density test 117-118

Neoprene 120, 122

Nomex 127, 128

Oil system design 76-77

Optical density. See Specific optical density.

Pain threshold 34-35

PBI(polybenzimidazole) 120, 127, 128, 129

Polycarbonate 120, 130-131

222

A: 222
"INDEX (Continued)

Polyethylene 109, 131

Polyimide 120, 126, 127

Polyisocyanurate foam 129, 132

Polymers, char-forminq 120, 129, 135

Polyphosphazene 120, 126

Polyurethane foam 106, 109, 111, 121-126

Pressure refueling 70

Quantity indicator, fuel 71, 73-74

Respiratory tract 36, 41

Restraint systems 103, 135, 137

Seat cushions 103, 135-136

flammability of, in large-scale tests 120

flammability test requirements for 137, 138

foams for, properties of 121-127

Self-sealing breakaway valves 66-69

design criteria for 56, 61, 68-69

for filters and strainers 74

for fuel lines 56, 60-61

for fuel tanks 67-68

Shielding of electrical compartments and wiring 83, 92, 95

to prevent flammable fluid flow 92-95

Smoke

effect of flame retardants on 111-112

generation of, by burning materials 110-111, 122

hazards of 30, 41

light obscuration by 110, 176-180

tests for 117-118, 136, 138-140

Sparks 80, 96-98

Specific optical density 110, 118

Sponsons 156-157

Structural panels, interior 129, 135, 137, 138-139

TB 746-93-2 181, 183

Temperatures, during postcrash fires 27-29

Thermoplastic materials, properties of 129-131 Toxic gases

from burning interior materials 113-114, 122, 127, 135 human tolerance to 37-41 , types generated during aircraft fires 31-33

Toxicity, animal tests for 114, 118-119

Transmitter 185, 186, 197-198

Transportation Systems Center (TSC) data bank 139

UMTA (Urban Mass Transportation Administration)

recommendations for flammability and smoke emission 116-117,

120, 140-141

Urethane. See Polyurethane foam.

S ( 223

Noma'.. .. .................................... .

• l i l I I I I4 • A •.
INDEX (Continued)

Velocity change

during crasiles 76, 188-190

from vibration 190-195

Vents, fuel 69-70

Vibration 188, 190-193

Visibility, underwater 149-150

U.S. CoVermnont Printing OWCOR~ 1970- 439-%24/1103 Segion Me. 3-11

224
DEPARTMENT OF THE ARMY

APPLIED TECHNOLOGY LABORATORY

US ARMY REStARCH 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)

FROM:

TO: Director, Applied Technology Laboratory, US Army Research and Technology

Laboratories (AVRADCOM), ATTN: DAVDL-ATL-ASV, Fort Eustis, Va. 23804

REMARKS:

SI GNATURE: _____________DATE: _______

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