NATIONAL

TRANSPORTATION

SAFETY BOARD

WASHINGTON, D.C. 20594

AIRCRAFT ACCIDENT REPORT

PB2000-910403

NTSB/AAR-00/03

DCA96MA070

In-flight Breakup Over the Atlantic Ocean

Trans World Airlines Flight 800

Boeing 747-131, N93119 

Near East Moriches, New York

July 17, 1996

6788G

Aircraft Accident Report

In-flight Breakup Over The Atlantic Ocean

Trans World Airlines Flight 800 

Boeing 747-131, N93119 

Near East Moriches, New York 

July 17, 1996

NTSB/AAR-00/03

DCA96MA070

PB2000-910403 National Transportation Safety Board

Notation 6788G 490 LíEnfant Plaza, S.W.

Adopted August 23, 2000 Washington, D.C. 20594
National Transportation Safety Board. 2000. In-flight Breakup Over The Atlantic Ocean, Trans World

Airlines Flight 800, Boeing 747-131, N93119, Near East Moriches, New York, July 17, 1996. Aircraft

Accident Report NTSB/AAR-00/03. Washington, DC.

Abstract: This report explains the accident involving Trans World Airlines, Inc. flight 800, which

experienced an in-flight breakup and then crashed into the Atlantic Ocean near East Moriches, New York,

on July 17, 1996. Safety issues in the report focus on fuel tank flammability, fuel tank ignition sources,

design and certification standards, and the maintenance and aging of aircraft systems. Safety

recommendations concerning these issues are addressed to the Federal Aviation Administration.

The National Transportation Safety Board is an independent Federal agency dedicated to promoting aviation, railroad, highway, marine,

pipeline, and hazardous materials safety. Established in 1967, the agency is mandated by Congress through the Independent Safety Board

Act of 1974 to investigate transportation accidents, determine the probable causes of the accidents, issue safety recommendations, study

transportation safety issues, and evaluate the safety effectiveness of government agencies involved in transportation. The Safety Board

makes public its actions and decisions through accident reports, safety studies, special investigation reports, safety recommendations, and

statistical reviews.

Recent publications are available in their entirety on the Web at <http://www.ntsb.gov>. Other information about available publications also

may be obtained from the Web site or by contacting:

National Transportation Safety Board

Public Inquiries Section, RE-51

490 L'Enfant Plaza, S.W.

Washington, D.C. 20594

(800) 877-6799 or (202) 314-6551

Safety Board publications may be purchased, by individual copy or by subscription, from the National Technical Information Service. To

purchase this publication, order report number PB2000-910403 from:

National Technical Information Service

5285 Port Royal Road

Springfield, Virginia 22161

(800) 553-6847 or (703) 605-6000

The Independent Safety Board Act, as codified at 49 U.S.C. Section 1154(b), precludes the admission into evidence or use of Board reports 

related to an incident or accident in a civil action for damages resulting from a matter mentioned in the report. 
iii Aircraft Accident Report

Contents

Abbreviations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xi

Executive Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xvi

1. Factual Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.1 History of Flight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.2 Injuries to Persons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.3 Damage to Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.4 Other Damage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.5 Personnel Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.5.1 The Captain (Left Front Seat) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.5.2 The Captain/Check Airman (Right Front Seat) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

1.5.3 The Flight Engineer (Right Aft/Flight Engineer’s Seat) . . . . . . . . . . . . . . . . . . . . . . 5

1.5.4 The Flight Engineer/Check Airman (Left Aft/Cockpit Jump Seat) . . . . . . . . . . . . . . 5

1.6 Airplane Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

1.6.1 Boeing 747—General Description/Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.6.1.1 747-100 Wing Center Section and Center Wing Fuel Tank Description . . . . . . 12

1.6.1.2 747-100 Air Conditioning Equipment Description . . . . . . . . . . . . . . . . . . . . . . . 16

1.6.1.3 747-100 Electrical and Wiring Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

1.6.1.3.1 747-100 Electrical Information. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

1.6.1.3.2 747 Wiring Information—General and Specific to Accident Airplane . . . . . 22

1.6.1.3.3 Separation of 747 Wire Circuits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

1.6.2 747-100 Fuel System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

1.6.2.1 Fuel Quantity Indication System Components and Wiring Information. . . . . . . 31

1.6.2.2 747-100 Fuel Pump System Description. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

1.6.2.3 747-100 Fuel Flow Indication Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

1.6.2.4 TWA Flight 800 Fueling Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

1.6.3 Maintenance Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

1.6.3.1 TWA’s 747 Maintenance Inspection Program . . . . . . . . . . . . . . . . . . . . . . . . . . 43

1.6.3.1.1 TWA’s General Inspection Policies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

1.6.3.1.2 TWA and Boeing Wiring Inspection Guidance . . . . . . . . . . . . . . . . . . . . . . . 46

1.6.3.2 Accident Airplane’s Maintenance Information. . . . . . . . . . . . . . . . . . . . . . . . . . 46

1.6.3.2.1 Airworthiness Directives and Service Bulletins Applicable 

to 747 Fuel Pumps and Related Wiring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

1.6.3.2.2 Airworthiness Directives and Service Bulletins Applicable 

to 747 Structural Inspections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

1.6.3.2.3 Maintenance Accomplished Near Fuel Quantity Indication System 

Wiring in the Accident Airplane. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

1.6.3.2.4 TWA Flight 800 Predeparture Maintenance Information . . . . . . . . . . . . . . . 54

1.7 Meteorological Information. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

1.8 Aids to Navigation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

1.9 Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

1.10 Airport Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

1.11 Flight Recorders. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58
Contents iv Aircraft Accident Report

1.11.1 Cockpit Voice Recorder. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

1.11.1.1 Sound Spectrum Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

1.11.1.2 Cockpit Voice Recorder-Related Airplane Tests . . . . . . . . . . . . . . . . . . . . . . . . 60

1.11.2 Flight Data Recorder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

1.12 Wreckage Recovery and Documentation Information . . . . . . . . . . . . . . . . . . . . . . . . . 62

1.12.1 Wreckage Recovered from the Red Zone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 69

1.12.2 Wreckage Recovered from the Yellow Zone. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

1.12.3 Wreckage Recovered from the Green Zone. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

1.12.4 Engines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74

1.12.5 Center Wing Fuel Tank Pumps and Other Components . . . . . . . . . . . . . . . . . . . . . 75

1.12.6 Air Conditioning Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

1.12.6.1 Brown Splatter Material on Air Conditioning Ducts . . . . . . . . . . . . . . . . . . . . . 78

1.12.7 Recovered Electrical Components/Wiring. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

1.13 Medical and Pathological Information. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84

1.14 Fire/Explosion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85

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Contents viii Aircraft Accident Report

Figures

1. Three views of the 747-100 airplane. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2. Location of some body stations and waterlines in the 747. . . . . . . . . . . . . . . . . . . . . . . . . . . 9

3a. Location of fuselage sections 41, 42, and 44 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3b. Location of fuselage sections 46 and 48 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

4a. A cross-section of the 747-100 wing center section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4b. Another cross-section of the 747-100 wing center section . . . . . . . . . . . . . . . . . . . . . . . . . . 14

5. The 747-100 aerodynamic fairings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

6. A schematic diagram of the 747-100ís air conditioning system . . . . . . . . . . . . . . . . . . . . . 18

7. Top and side views of the 747-100ís wing center section and air conditioning system . . . . 19

8. Underside view of the 747-100ís wing center section and air conditioning system . . . . . . . 20

9. The air conditioning ducts around the 747-100ís center wing fuel tank . . . . . . . . . . . . . . . 21

10. Diagram of a BMS13-42A Poly-X wire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

11. A schematic diagram of the 747-100 fuel tank arrangement . . . . . . . . . . . . . . . . . . . . . . . . 27

12. A layout of a typical 747-100 center wing fuel tank and left wing fuel tank vent system . . 29

13. A 747 pressure fueling flow diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

14. A diagram of the 747 pressure fueling system components . . . . . . . . . . . . . . . . . . . . . . . . . 32

15. Diagrams of a fuel quntity probe and a compensator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

16. The fuel quantity probe and compensator locations in the 747ís center wing fuel tank . . . . 36

17. A schematic diagram of the 747-100ís fuel quantity indication system wiring, 

fuel flow wiring, cockpit voice recorder wiring, and other wiring information . . . . . . . . . . 38

18. Locations of the wire raceways in the 747-100. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

19. A schematic diagram of the fuel pump locations in the 747 . . . . . . . . . . . . . . . . . . . . . . . . . 41

20. A diagram of the 747-100 interior. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

21. The wreckage location relative to the airplaineís flightpath, JFK, and Long Island. . . . . . . 64

22a. Map showing the locations of the red, yellow, and green zones. . . . . . . . . . . . . . . . . . . . . . 66

22b. A 747, color-coded to indicate the debris fields from which corresponding 

wreckage was recovered . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

22c. The accident airplaneís flightpath, color-coded to indicate the debris fields from 

which corresponding wreckage was recovered. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68

23. A diagram of a jettison/override fuel pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

24. A photograph showing wreckage recovered with tangled and damaged wires attached . . . 80

25. Radar data showing vehicle and/or object tracks within 10 nm of TWA flight 800 

just before the accident . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

26. Primary radar returns that appeared near the accident airplane after 2031:12 . . . . . . . . . . . 91
Contents ix Aircraft Accident Report

27. Sequence of eight primary radar returns recorded by the Islip, New York, radar 

site about the time of the accident . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92

28a. Roll angle for nose off at 2031:15.2 cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98

28b. Map view of nose off at 2031:15.2 cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98

28c. Altitude for nose off at 2031:15.2 cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99

28d. East position for nose off at 2031:15.2 cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99

28e. North position for nose off at 2031:15.2 cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100

29. A photograph of the right side of the large three-dimensional reconstruction, 

with the support scaffolding visible . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102

30. A schematic drawing of the lower portion of the fuselage skin in the area 

forward of the front spar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105

31. Diagram showing the approximate shape of the upper skin panel stiffener and 

its location before being struck by SWB3 as the beam rotated forward . . . . . . . . . . . . . . . 107

32. A flow chart of the fuel-related research conducted to support this investigation . . . . . . . 121

33. Side view, top view, and three-dimensional view of wing center section/center 

wing fuel tank bays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127

34. A trip-curve chart for a typical circuit breaker (excerpted from Texas 

Instruments KlixonÆ specification) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 149

35a. Sample origination locations for the 747 manufactured in 1970 . . . . . . . . . . . . . . . . . . . . 152

35b. Sample origination locations for the DC-10 manufactured in 1973 . . . . . . . . . . . . . . . . . . 152

35c. Sample origination locations for the 747 manufactured in 1973 . . . . . . . . . . . . . . . . . . . . 153

36. Chart showing the percent probability of failure of Kapton wiring in different 

locations in the P-3 airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 158

37. A diagram of a scavenge pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171

38. Diagram showing the areas typically inspected on 747 airplanes. . . . . . . . . . . . . . . . . . . . 190

39. A chart showing the 737 fuel pump conduit wear rate that was identified by 

the FAA during representative 500 wire inspections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214

40. Map showing the witness locations and the location of the accident site. . . . . . . . . . . . . . 231

41. A map showing the debris zones and the positions of the five witnesses of 

special interest. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 244

42. Map of sight lines plotted for 11 witnesses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 249

43. Diagram of warhead detonation fragment pattern. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 253

44. Diagram showing airplane and missile launch site locations that could result in missile

 selfdestruct within 40 feet of the plane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254

45. A graphical illustration of the equidistance tendency . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 268

46. Diagram of the center wing tank fuel quantity indication system wiring on the 747 

and its common connectors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 282

47. A time line of salvage operation events . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366

Sooting and fracture diagrams. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 378-411
Contents x Aircraft Accident Report

Tables

1. Injury chart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2. TWAís fueling records, indicating TWA flight 800ís fuel loading 

when it prepared to depart JFK . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

3. Winds aloft measured by a weather balloon launched 

from Upton, New York . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

4. A time line of simulation scenarios. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 97

5. 747-100 wing center section/center wing fuel tank 

SWBs and spars failure strengths . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120

6. A summary of the maximum and minimum recorded ullage temperatures 

at 13,700 feet msl during the emulation flight test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 128

7. Fuel/air mass ratios and ullage temperatures measured at the

bay 2 sampling port during the TWA flight emulation flight test. . . . . . . . . . . . . . . . . . . . 130

8. Results of wet short-circuit tests conducted by Lectromec. . . . . . . . . . . . . . . . . . . . . . . . . 146

9. The origin and potential environmental exposures of test samples. . . . . . . . . . . . . . . . . . . 151

10. Several fault tree analysis basic elements, as identified in Boeingís 

fault tree analysis report, with their failure rates, exposure times, and 

calculated event probability. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 166

11. Some of the basic elements of Boeingís fault tree analysis, showing 

the differences between the data presented in Boeingís graphical and 

tabular representations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167
xi Aircraft Accident Report

Abbreviations

AAIB Air Accidents Investigation Branch (British)

a.c. alternating current

AC advisory circular

ACAP Aviation Consumer Action Project

ACM air cycle machine

AD airworthiness directive

AFCB arc-fault circuit breaker

AFM aircraft flight manual

AFRL U.S. Air Force Research Laboratory 

AIDS airborne integrated data system

ALPA Air Line Pilots Association

amp ampere

AOA angle-of-attack

API (2003) American Petroleum Institute (Recommended Practice 2003)

APU auxiliary power unit

ARAC Aviation Rulemaking Advisory Committee

ARTCC air route traffic control center

AS aircraft service (check)

ASB alert service bulletin

ASR airport surveillance radar

ATA Air Transport Association of America

ATC air traffic control

ATE automated test equipment

ATF Bureau of Alcohol, Tobacco, and Firearms

ATP airline transport pilot

ATSRAC Aging Transport Systems Rulemaking Advisory Committee

AWIGG Aircraft Wiring and Inert Gas Generator 

BL butt line

BMS Boeing Material Specification

CAM cockpit area microphone

CAMI Civil Aeromedical Institute 

(Federal Aviation Administration)
Abbreviations xii Aircraft Accident Report

CDG Charles DeGaulle International Airport (Paris, France)

CDL Combustion Dynamics Limited

CEU central electronics unit

CFR Code of Federal Regulations

c.g. center of gravity

CHDO Certificate Holding District Office

CIA Central Intelligence Agency

CIT California Institute of Technology

CMR Christian Michelsen Research

CVR cockpit voice recorder

CWT center wing fuel tank

d.c. direct current

DERA Defense Evaluation and Research Agency (British)

DGAC Direction Generale de LíAviation Civile

DNA deoxyribonucleic acid 

DoD Department of Defense

DRI Desert Research Institute

EDS energy dispersive spectroscopy

E/E electrical/electronics

EME electromagnetic environment

EMI electromagnetic interference

EPR engine pressure ratio

F Fahrenheit

FAA Federal Aviation Administration

FARs Federal Aviation Regulations

FBI Federal Bureau of Investigation

FDR flight data recorder

FMEA failure modes and effects analysis

FOK Francis S. Gabreski Airport 

(Westhampton Beach, New York)

fps feet per second

FQIS fuel quantity indication system

FR Federal Register

FTHWG Fuel Tank Harmonization Working Group
Abbreviation xiii Aircraft Accident Report

GPS global positioning system

HF high frequency

Hg mercury

HIRF high-intensity radiated fields

HP High Performance

HPC high-pressure compressor

HSSM Hamilton Standard Service Manual

Hz hertz (cycles per second)

IAM International Association of Machinists, Aerospace Workers, 

and Flight Attendants

IOE initial operating experience

J joule

JAA Joint Aviation Authorities

JFK John F. Kennedy International Airport (New York, New York)

JSC Joint Spectrum Center (DoD)

kJ kilojoule

kW kilowatt

LaRC Langley Research Center

LBL left butt line

LFL lower flammability limit

LLS laser line-scanning

LPC low-pressure compressor

MAC mean aerodynamic chord

MHz megahertz

MIE minimum ignition energy

mil one-thousandth of an inch

mJ millijoule

mm millimeter

MO modification order

MSFC George C. Marshall Space Flight Center 

MSIC Missile and Space Intelligence Center 

(Defense Intelligence Agency)
Abbreviations xiv Aircraft Accident Report

msl mean sea level

NaCl saline-water electrolyte solution

NAFI U.S. Naval Avionics Facility

NASA National Aeronautics and Space Administration

NASIP National Aviation Safety Inspection Program

NAWC-AD Naval Air Warfare CenteróAircraft Division

NAWC-WD Naval Air Warfare CenteróWeapons Division

nm nautical mile

NOAA National Oceanic and Atmospheric Administration

NPRM notice of proposed rulemaking

NRC Nuclear Regulatory Commission

NRL Naval Research Laboratory

NYANG New York Air National Guard

PED personal electronic device

PETN pentaerythritol tetranitrate

PI production illustration (Boeingís)

PMI principal maintenance inspector

P/N part number

PPM partial program manager

PS periodic service (check)

psi (pressure expressed in) pounds per square inch

PTFE polytetrafluoroethylene

P&W Pratt & Whitney

RADES Radar Evaluation Squadron

RBL right butt line

RDX cyclotrimethylenetrinitramine

RF radio frequency

ROV remote-operated vehicle

SAE Society of Automotive Engineers

SB service bulletin

SDR service difficulty report

SEM scanning electron microscope

SFAR Special Federal Aviation Regulation

SL service letter
Abbreviations xv Aircraft Accident Report

S/N serial number

SNL Sandia National Laboratories

SS station service (check)

SSS side-scan sonar

STA body station 

STC supplemental type certificate

STP surge tank protection

SUPSALV Supervisor of Salvage and Diving

SURVIAC Survivability/Vulnerability Information Analysis Center

SWB spanwise beam

SWPM Standard Wiring Practices Manual (Boeing)

TC type certificate

TCS time control service (check)

TDR time-domain reflectometry

TRU transformer-rectifier unit

TSD transient suppression device

TWA Trans World Airlines, Inc.

UDRI University of Dayton Research Institute

UL Underwriters Laboratories, Inc.

USAF U.S. Air Force

USCG U.S. Coast Guard

UV ultraviolet ray

VHF very high frequency

V/m volt/meter

VSCU video system control unit

VSO volumetric shutoff

WCS wing center section

WHCASS White House Commission on Aviation Safety and Security

WL body waterline

WS wing station

WSSIWG Wire System Safety Interagency Working Group

XL-ETFE cross-linked ethylenetetrafluoroethylene
xvi Aircraft Accident Report

Executive Summary

On July 17, 1996, about 2031 eastern daylight time, Trans World Airlines, Inc.

(TWA) flight 800, a Boeing 747-131, N93119, crashed in the Atlantic Ocean near East

Moriches, New York. TWA flight 800 was operating under the provisions of 14 Code of

Federal Regulations Part 121 as a scheduled international passenger flight from John F.

Kennedy International Airport (JFK), New York, New York, to Charles DeGaulle

International Airport, Paris, France. The flight departed JFK about 2019, with 2 pilots,

2 flight engineers, 14 flight attendants, and 212 passengers on board. All 230 people on

board were killed, and the airplane was destroyed. Visual meteorological conditions

prevailed for the flight, which operated on an instrument flight rules flight plan.

The National Transportation Safety Board determines that the probable cause of

the TWA flight 800 accident was an explosion of the center wing fuel tank (CWT),

resulting from ignition of the flammable fuel/air mixture in the tank. The source of

ignition energy for the explosion could not be determined with certainty, but, of the

sources evaluated by the investigation, the most likely was a short circuit outside of the

CWT that allowed excessive voltage to enter it through electrical wiring associated with

the fuel quantity indication system. 

Contributing factors to the accident were the design and certification concept that

fuel tank explosions could be prevented solely by precluding all ignition sources and the

design and certification of the Boeing 747 with heat sources located beneath the CWT

with no means to reduce the heat transferred into the CWT or to render the fuel vapor in

the tank nonflammable.

The safety issues in this report focus on fuel tank flammability, fuel tank ignition

sources, design and certification standards, and the maintenance and aging of aircraft

systems. Safety recommendations concerning these issues are addressed to the Federal

Aviation Administration.
1 Aircraft Accident Report

1. Factual Information

1.1 History of Flight

On July 17, 1996, about 2031 eastern daylight time,1 Trans World Airlines, Inc.

(TWA) flight 800, a Boeing 747-131 (747), N93119, crashed in the Atlantic Ocean near

East Moriches, New York. TWA flight 800 was operating under the provisions of 14 Code

of Federal Regulations (CFR) Part 121 as a scheduled international passenger flight from

John F. Kennedy International Airport (JFK), New York, New York, to Charles DeGaulle

International Airport (CDG), Paris, France. The flight departed JFK about 2019, with

2 pilots, 2 flight engineers, 14 flight attendants, and 212 passengers on board. All

230 people on board were killed, and the airplane was destroyed. Visual meteorological

conditions prevailed for the flight, which operated on an instrument flight rules flight plan.

On the day of the accident, the accident airplane departed Athens, Greece, as TWA

flight 881 about 0537, landed at JFK about 1631, and arrived at terminal 5, gate 27,

about 1638. The flight crew that had flown the accident airplane from Athens, Greece, to

JFK told National Transportation Safety Board investigators that it did not observe any

operational abnormalities during that flight. A scheduled flight crew change occurred at

JFK. The accident airplane was refueled at JFK2 and remained at gate 27 with the

auxiliary power unit (APU) and two of its three air conditioning packs operating (for

about 2 1/2 hours) until it departed as TWA flight 800. 

According to company records, flight 800 was to be the first leg of a scheduled

3-day trip sequence for the four flight crewmembers. A captain occupied the left front

seat, a captain/check airman occupied the right front seat, a flight engineer occupied the

right aft seat (flight engineer position), and a flight engineer/check airman occupied the

left aft seat (cockpit jump seat).3 TWA flight 800 was scheduled to depart JFK for CDG

about 1900; however, the flight was delayed because of a disabled piece of ground

equipment4

 and concerns about a suspected passenger/baggage mismatch. According to

the cockpit voice recorder (CVR),5 at 1959:44, gate agent personnel advised the flight

1 Unless otherwise indicated, all times are eastern daylight time, based on a 24-hour clock. 

2 For additional information regarding TWA flight 881 flight crew statements about the airplaneís fuel

load during that flight, refueling difficulties at JFK, and the fuel on board the accident airplane when it

departed JFK, see section 1.6.

3 The flight engineer was receiving initial operating experience (IOE) training during the accident

flight. 

4 A disabled fleet service vehicle was blocking the accident airplane at the gate, and there was a delay

obtaining the proper equipment to tow the fleet service vehicle. 

5 See appendix B for a complete transcript of the CVR. The local (eastern daylight) time reference was

established by correlating the last seven very high frequency (VHF) radio transmissions made by the pilots

of TWA flight 800 and recorded by the flight data recorder (FDR) microphone keying parameter with the

CVR and the Boston Air Route Traffic Control Center (ARTCC). The correlation points all agreed within

1 second. 
Factual Information 2 Aircraft Accident Report

crew that, although a passengerís bag had been pulled because they suspected that it was

unattended, they subsequently confirmed that ìthe passenger was on board the whole

time.î The CVR recorded the sound of the cockpit door closing at 1959:59, and the flight

crew of TWA flight 800 continued to prepare for departure. 

TWA flight 800 was pushed back from the gate about 2002. According to the

CVR, between 2005 and 2007:46, the flight crew started the Nos. 1, 2, and 4 engines and

completed the after-start checklist. At 2007:52, the captain/check airman advised the JFK

gate hold controller that TWA flight 800 was ìready to taxi.î About 2008, the flight crew

received taxi instructions from air traffic control (ATC) and began to taxi to runway 22

right (22R), the departure runway. While the airplane was taxiing (about 2014), the CVR

recorded the flight crew starting the No. 3 engine and conducting the delayed engine-start

and taxi checklists.

At 2017:18, the CVR recorded ATC stating, ìTWA 800 heavy caution wake

turbulence from a 757, runway 22R, taxi into position and hold.î The CVR transcript

indicated that the captain/check airman acknowledged the ATC clearance and that the

airplane was taxied into position on the departure runway.6

 At 2018:21, ATC advised the

pilots of TWA flight 800 that the wind was out of 240° at 8 knots and cleared flight 800 for

takeoff on runway 22R. The CVR recorded the flight crew conducting the before-takeoff

checklist and the sound of increasing engine noise; FDR and CVR information indicated

that the airplane became airborne about 2019.

During the accident airplaneís departure from JFK, the pilots of TWA flight 800

received a series of (generally increasing) altitude assignments and heading changes from

New York Terminal Radar Approach Control and Boston ARTCC controllers. At 2025:41,

Boston ARTCC advised the pilots to ìclimb and maintain [19,000 feet] and expedite

through [15,000 feet]î; the pilots acknowledged the instructions at 2025:47. According to

the CVR, at 2026:24, Boston ARTCC amended TWA flight 800ís altitude clearance,

advising the pilots to maintain 13,000 feet mean sea level (msl). At 2027:47, the CVR

recorded the sound of the altitude alert tone, and the FDR data indicated that the airplane

reached its assigned altitude.

At 2029:15, the CVR recorded the captain stating, ìLook at that crazy fuel flow

indicator there on number fourÖsee that?î7

 At 2030:15, Boston ARTCC advised,

ìTWA 800 climb and maintain [15,000 feet msl].î The CVR recorded the captain stating,

ìclimb thrust,î and the captain/check airman acknowledging the ATC clearance at

2030:18. At 2030:25, the captain repeated, ìclimb thrust,î and at 2030:35, the flight

engineer responded, ìpowerís set.î The CVR recording of the next 30 seconds from the

cockpit area microphone (CAM)8 includes the following sounds:

6 Review of the ATC and CVR transcripts from the accident flight indicated that the captain (left front

seat) was performing the pilot-flying duties, while the captain/check airman (right front seat, identified as

the ìfirst officerî in the CVR transcript) was performing the radio communications and other pilot-notflying duties during the departure. 

7 For additional information regarding the fuel flow sensing and indication system on the 747-100, see

section 1.6.2.3. 
Factual Information 3 Aircraft Accident Report

ï a sound similar to a mechanical movement in the cockpit (at 2030:42), 

ï an unintelligible word (at 2031:03), and 

ï sounds similar to recording tape damage noise (at 2031:05).9

At 2031:12, the CVR recording ended. A sound spectrum study of the information

recorded by the CVR revealed that twice within the last second of the CVR recording

(about 0.73 and 0.68 seconds before the recording stopped), the captainís channel

recorded harmonic tones at the 400 Hertz10 (Hz) frequency, but it did not record other

electrical system background noise that it had recorded previously throughout the

recording. These other electrical system background noises were recorded on the other

CVR channels without interruption.11 The CVR then recorded a ìvery loud soundî for a

fraction of a second (0.117 second) on all channels immediately before the recording

ended. The accident airplaneís last recorded radar12 transponder return occurred at

2031:12, and a review of the FDR data indicated that the FDR lost power at 2031:12. 

According to the Boston ARTCC transcript, at 2031:50, the captain of an Eastwind

Airlines Boeing 737 (Stinger Bee flight 507) reported that he ìjust saw an explosion out

here.î About 10 seconds later, the captain of Stinger Bee flight 507 further advised, ìwe

just saw an explosion up ahead of us hereÖabout 16,000 feet or something like that, it just

went down into the water.î13 Subsequently, many ATC facilities in the New York/Long

Island area received reports of an explosion from other pilots operating in the area.

Many witnesses in the vicinity of the accident at the time that it occurred stated

that they saw and/or heard explosions, accompanied by a large fireball over the ocean, and

observed debris, some of which was burning, falling to the water. According to witness

documents, about one-third of these witnesses reported that they observed a streak of

light, resembling a flare, moving upward in the sky to the point where a large fireball

appeared. Several witnesses reported seeing this fireball split into two fireballs as it

descended toward the water.14

8 The audio information recorded by the CAM includes audio (airborne) sounds and vibrations picked

up by the fuselage-mounted microphone. 

9 Subsequent examination of the CVR tape indicated that these sounds were likely the result of water

damage to the tape head. Because of the position of the tape on the reels, the portion of the tape that

contained sounds from the last seconds before the CVR stopped recording was exposed to water after the

accident. 

10 A Hz is a unit of frequency equal to 1 cycle per second.

11 The Safety Boardís CVR sound spectrum study revealed no other such changes or other unusual

electrical occurrences on the 31-minute 30-second recording. See section 1.11.1.1 for a detailed discussion

of the results of the sound spectrum study and other CVR-related testing, including how the CVR system

installed on the 747 would respond to various types of explosive events. 

12 Examination of radar data indicated that the accident airplaneís last radar transponder return was

recorded by the Federal Aviation Administrationís (FAA) radar site at Trevose, Pennsylvania. For additional

information regarding radar information, see section 1.16.1. 

13 Radar data and ATC records indicated that Stinger Bee flight 507 was about 20 to 25 miles northeast

of TWA flight 800, on a southwesterly heading. 

14 For additional information regarding witness reports, including that of the pilot of Stinger Bee flight

507, see section 1.18.4.
Factual Information 4 Aircraft Accident Report

Pieces of the airplane wreckage were discovered floating on and beneath the

surface of the Atlantic Ocean about 8 miles south of East Moriches, New York. The main

wreckage was found on the ocean floor, between 40° 37' 42" and 40° 40' 12" north latitude

and 72° 40' 48" and 72° 35' 38" west longitude. The accident occurred in dusk lighting

conditions. 

1.2 Injuries to Persons

1.3 Damage to Aircraft

The airplane was destroyed by the explosion, breakup and impact forces, and fire.

According to insurance company records, the airplane was valued at $11 million. 

1.4 Other Damage

No structures on the ground were damaged. 

1.5 Personnel Information

The Safety Board reviewed the flight crewís flight- and duty-time limits and rest

records and found no evidence that they were not within the limits established by Federal

regulations. The cabin crew comprised 14 flight attendants. 

1.5.1 The Captain (Left Front Seat)

The captain, age 58, was hired by TWA on May 20, 1965. He held airline transport

pilot (ATP) certificate No. 1453736 with multiengine land and instrument ratings. The

captainís most recent FAA first-class airman medical certificate was issued on March 22,

1996, and contained the limitation that he possess glasses for near vision. 

Table 1. Injury chart.

Injuries Flight Crew Cabin Crew Passengers Other Total

Fatal 4 14 212 0 230

Serious 0 0 0 00

Minor 0 0 0 00

None 0 0 0 00

Total 4 14 212 0 230
Factual Information 5 Aircraft Accident Report

The captainís first assignment with TWA was as a first officer on the Convair 880.

He subsequently served as first officer and captain on the Boeing 707 (707) and

Lockheed 1011 before transitioning to the 747. He received a 747 type rating on

February 19, 1990, and began 747 captain upgrade training on May 21, 1996; the captainís

most recent proficiency check was satisfactorily completed on June 19, 1996. According

to TWA records, at the time of the accident, the captain had flown approximately 18,800

total flight hours, including 5,490 hours in the 747. 

1.5.2 The Captain/Check Airman (Right Front Seat)

The captain/check airman, age 57, was hired by TWA on April 13, 1964. He held

ATP certificate No. 1475512 with single-engine land, multiengine land, and instrument

ratings. The captain/check airmanís most recent FAA first-class airman medical certificate

was issued on April 15, 1996, with no restrictions or limitations.

The captain/check airmanís first assignment with TWA was as a first officer on the

Convair 880. He subsequently served as flight crewmember on the 707, Boeing 727, and

Lockheed 1011 before transitioning to the 747. He received a 747 type rating on

December 30, 1974, and qualified as a 747 check airman on May 3, 1993. According to

TWA records, at the time of the accident, the captain/check airman had flown

approximately 17,000 total flight hours, including 4,700 hours in the 747. 

1.5.3 The Flight Engineer (Right Aft/Flight Engineerís Seat)

The flight engineer trainee, age 24, was hired by TWA on June 22, 1996. He held

flight engineer-turbo jet-powered certificate No. 306804492. The flight engineer traineeís

most recent FAA first-class airman medical certificate was issued on December 12, 1995,

with no restrictions or limitations.

The flight engineer trainee was hired by TWA as a 747 flight engineer, and the

accident flight was the sixth leg of his IOE training. According to TWA records, at the

time of the accident, the flight engineer trainee had flown approximately 2,520 total flight

hours, including about 30 hours as flight engineer trainee in the 747. 

1.5.4 The Flight Engineer/Check Airman (Left Aft/Cockpit Jump 

Seat) 

The flight engineer/check airman, age 62, was hired by TWA on February 26,

1966. He held ATP certificate No. 1409009 and flight engineer-turbo jet-powered

certificate No. 1694661. The flight engineer/check airmanís most recent FAA first-class

airman medical certificate was issued on July 17, 1996, and contained the limitation that

he wear glasses for near and distant vision. 

The flight engineer/check airmanís first assignment with TWA was as a flight

engineer on the Lockheed Constellation 749, 749A, and 1049G model airplanes. In
Factual Information 6 Aircraft Accident Report

February 1967, he upgraded to the first officer position on these model airplanes. On

November 19, 1986, the flight engineer/check airman received a 747 type rating. He

subsequently performed 747 first officer and captain duties until he turned 60 on July 2,

1993, at which time, he became a flight engineer on the 747 rather than retire.15 According

to TWA records, at the time of the accident, the flight engineer/check airman had about

3,047 hours of flight engineer experience, including 2,397 hours as flight engineer on the

747.

1.6 Airplane Information

The accident airplane, N93119, a 747-100 series airplane (model 747-131),16 serial

number (S/N) 20083, was manufactured by Boeing in July 1971 and purchased new by

TWA. The airplane was added to TWAís operating certificate on October 27, 1971, and,

except for a 1-year period,17 was operated by TWA in commercial transport service until

the accident occurred. According to TWA records, the accident airplane had 93,303 total

hours of operation (16,869 flight cycles)18 at the time of the accident.19 The 747-100 is a

low-wing, transport-category airplane that is about 225 feet long and 63 feet high (from

the ground to the top of the vertical stabilizer), with a wingspan of about 195 feet. The

747-100 can carry about 430 passengers and cargo. Figure 1 shows three views of the

747-100 airplane. 

The accident airplane was equipped with four Pratt & Whitney (P&W) JT9D-7AH

turbofan engines. Company maintenance records indicated that the No. 1 (outboard left)

engine, S/N 662209, was installed on the accident airplane on December 31, 1995, and

had operated about 47,989 hours since new; the No. 2 (inboard left) engine, S/N 662593,

was installed on the accident airplane on December 6, 1995, and had operated about

80,884 hours since new; the No. 3 (inboard right) engine, S/N 662426, was installed on

the accident airplane on June 18, 1996, and had operated about 80,336 hours since new;

15 Federal Aviation Regulations (FARs) stipulate a mandatory retirement age of 60 for air carrier

captains and first officers; however, they allow continued airman activity in other areas, including the flight

engineer position. 

16 The 747-100 series airplane is one of several 747 models. Other 747 models include the -200, -300,

-SP, and -SR (collectively referred to as the ìClassicî) series airplanes, and the -400. The military uses

derivatives of the 747, which are identified as the E-4B and VC-25. 

17 On December 15, 1975, the accident airplane was ferried to the Boeing Military Aircraft Company in

Wichita, Kansas, to be prepared for sale to the Government of Iran. According to the Airclaims Limited

database, the accident airplane was sold to the Iranian Air Force in December 1975; however, the airplane

was not delivered to Iran and was returned to TWAís operating certificate on December 16, 1976.

Maintenance records indicated that the airplane had been operated about 7 hours with four landings between

December 1975 and December 1976. (See Maintenance Records Group Chairmanís Factual Report, dated

October 8, 1997, for a list of the modifications and inspections accomplished in preparation for the sale of

the airplane.)

18 A flight cycle is one complete takeoff and landing sequence.

19 According to a Boeing AERO magazine article, the minimum design service objective for 747 series

airplanes was 20,000 flights, 60,000 hours, and 20 years. 
Factual Information 7 Aircraft Accident Report

and the No. 4 (outboard right) engine, S/N 662463, was installed on the accident airplane

on May 11, 1996, and had operated about 77,061 hours since new.

According to TWAís dispatch documents for the accident flight, the airplaneís

takeoff weight was calculated to be 590,441 pounds,20 including 19,751 pounds of cargo

(6,062 pounds of cargo in the forward cargo compartment and 13,689 pounds of cargo in

the aft and bulk cargo compartments)21 and 176,600 pounds of fuel. TWA dispatch records

and load information recorded by the CVR indicated that there were 29 passengers in the

first-class cabin, 183 passengers in the coach cabin, and 18 crewmembers (4 flight

crewmembers and 14 cabin crewmembers) on board the airplane. The dispatch documents

indicated that the airplaneís takeoff center of gravity (c.g.) was calculated to be 18.4

percent of the mean aerodynamic chord (MAC), and the takeoff horizontal stabilizer trim

setting was 6.1 units nose up. 

1.6.1 Boeing 747óGeneral Description/Information

The 747-100 fuselage comprises five major sections (referred to in Boeing

manuals as sections 41 [the forward section], 42, 44, 46, and 48 [the tail section]) and

consists of the external skin, internal circumferential frames, and longitudinal (fore-aft)

stiffening members, called stringers.22 For ease of reference, in the airplaneís

documentation, Boeing divided the airplane into reference planes, which are designated as

body stations (STA),23 body waterlines (WL),24 and butt lines (BL).25 The locations of

various components and fuselage areas in 747s are measured in inches from fixed points

of reference. Figure 2 shows the locations of some STAs and WLs. Figures 3a and 3b

show the locations of numerous STAs and other fuselage areas, including fuselage

sections, doors, and wheel wells. 

20 The maximum certificated takeoff gross weight for the accident airplane was 734,000 pounds. 

21 According to Boeing documents, the aft cargo compartment consists of an aft containerized

cargo/baggage compartment and a (farther aft) bulk cargo compartment, which have separate doors and are

separated by compartment divider nets. TWAís dispatch records use the term ìrear cargo compartmentî to

describe the aft containerized cargo/baggage compartment and ìaft cargo compartmentî to describe the bulk

cargo compartment. These dispatch documents specified that the ìrear cargo compartmentî contained

12,428 pounds of cargo and that the ìaft cargo compartmentî contained 1,261 pounds of cargo. 

22 The 747-100 wings also contain stringers, which extend from the wing root to the wing tip, to support

and reinforce the wing skin. 

23 A STA is a longitudinal point along an airplaneís fuselage, identified numerically by its distance in

inches from a reference point. In a classic 747, this point is 90 inches forward of the airplaneís nose. 

24 A WL is a horizontal point along an airplaneís fuselage, identified numerically by its distance in

inches from a parallel imaginary plane (WL 0.00) located 91 inches below the lowest body surface. 

25 A BL number refers to a lateral fuselage location and reflects the number of inches the location is

right or left of the airplaneís centerline. 
Figure 1. Three views of the 747-100 airplane. 32’ 25’6” 78’11” 225’2” 10’ 231’4” 69’6” 2° TYP 12’6” 36’2” 72’9” 195’8” 
Figure 2. Location of some body stations and waterlines in the 747.
Figure 3a. Location of fuselage sections 41, 42, and 44.
Figure 3b. Location of fuselage sections 46 and 48. STA 1480 STA 1640 STA 1694 STA 1800 STA 1980 STA 1920 STA 2240 STA 2261 STA 2280STA 2360 STA 2360 STA 2484STA 2517 STA 2598 STA 2775

STABILHINGE 

STA 2658 STA 2638 S

2

STA 2594 ACCESS DOOR STA 2436 STA 2412 SECTION 48 SECTION 46 STA 2007 STA 1865 STA 1700 STA 1718 STA 1664 AFT CARGO DOOR BULK CARGO DOOR 

STA 2040
Factual Information 12 Aircraft Accident Report

 The vertical and horizontal stabilizers intersect the tail section of the fuselage

between STAs 2220 and 2590. The wing leading edges intersect the fuselage (with a

sweepback angle of 37 1/2o) just aft of STA 900, and the trailing edge is located just

forward of STA 1480. The wing front, mid, and rear spars extend through the wing center

section (WCS)26 between the right and left wings at STAs 1000, 1140.5, and 1241,

respectively (see figures 3a, 4a, and 4b). The fuselage is nearly circular in cross-section

where it intersects the front spar at the ring chord.27 The bottom of this portion of the

fuselage is also attached to the lower end of the keel beam (see figure 4a), which is a

box-shaped, load-bearing structure located along the airplaneís centerline that extends

from the aft wall of the forward cargo compartment28 (STA 1000) through the main and

body landing gear compartments to the forward wall of the aft cargo compartment

(STA 1480). (Figure 4a shows the keel beam structure and locations of WCS lateral

beams/spars, and the ring chord.) The portion of the airplane below the WCS is contained

within an aerodynamic fairing29 that blends into the wing leading edge. Figure 5 shows the

747-100 aerodynamic fairings. 

1.6.1.1 747-100 Wing Center Section and Center Wing Fuel Tank Description

The 747-100ís WCS is located aft of the forward cargo compartment and forward

of the main landing gear bay in the lower fuselage. Like the cargo compartments, the WCS

is below the main cabin floor. The WCS is about 21 feet wide, 20 feet long, and varies in

height from about 4 1/2 to 6 feet (with the shortest height located at the aft end). The WCS

is framed by the front and rear wing spars and the side-of-body ribs that separate the WCS

from the inboard wing fuel tanks. The upper and lower skin panels separate the WCS from

the passenger cabin floor (which is located above the WCS and supported by longitudinal

floor beams) and the airplaneís heat exchanger/air conditioning equipment (which is

located beneath the WCS),30 respectively. According to Boeing, the WCS carries the wing

bending forces through the airplane and supports the fuselage during flight.

26 The WCS is a large structural box located in the 747 lower fuselage between the wings and comprises

the center wing fuel tank (CWT) and a dry bay directly forward of the CWT. For additional information

regarding the WCS and CWT, see section 1.6.1.1. (Note: The CWT is called the wing center tank in some

documents.) 

27 A chord is an angle member that attaches two other pieces together, usually at an angle of about 90o.

The ring chord is an angle member that attaches the bottom of the forward fuselage section to the front side

of the lower-pressure bulkhead and the front spar. 

28 As previously discussed, the 747 has a forward and an aft cargo compartment, each located in the

lower portion of the fuselage, below the main cabin floor.

29 The fairings are composite material shaped around the fuselage and wings to make the airplane more

aerodynamic. 

30 For further details about the airplaneís heat exchanger/air conditioning equipment (air conditioning

packs), see section 1.6.1.2.
Factual Information 13 Aircraft Accident Report

Figure 4a. A cross-section of the 747-100 wing center section. 

Ring Chord

Front Spar

Vent

Opening

Vent

Opening

Forward

Keel Beam

Spanwise Beam 3

Intercoastal

Spanwise Beam 2

Midspar

Spanwise

Beam 1

Rear Spar

Floor Beams
Figure 4b. Another cross-section of the 747-100 wing center section.
Figure 5. The 747-100 aerodynamic fairings.
Factual Information 16 Aircraft Accident Report

The WCS is divided laterally into compartments by six beams, including (from

farthest aft to forward) the rear spar, spanwise beam (SWB)1, the mid spar, SWB2,

SWB3, and the front spar (see figures 4a and 4b). In addition, a partial longitudinal rib

divides the WCS along the airplaneís centerline (BL 0) between the mid and rear spars.

The WCS compartments located between the rear spar and SWB3 comprise the CWT. In

the 747-100, the compartment between SWB3 and the front spar is a dry bay that is not

intended to contain fuel.31 SWBs 1 and 2, the partial ribs, and the mid spar in the CWT

contain a number of cutouts and vent holes to allow air or fuel to move between the

various bays and electrical tubing/conduit32 to pass through. In addition, the SWBs and

mid spar contain maintenance access doors, which are attached with removable fasteners.

SWB2 also contains a manufacturing access door, which was permanently fastened in

place during completion of the manufacturing process. The CWT is sealed from the

outside atmosphere except for two vent stringers that connect the CWT to vents in the

surge tanks at each of the two wing tips. The WCS dry bay is vented to the atmosphere

through two openings in the front spar (see figure 4a). 

Attached to the bottom of the WCS is the keel beam (see figures 4a and 10), which

provides longitudinal strength to the airplane. The keel beam contains two vertical webs

(located at left BL [LBL] 9 and right BL [RBL] 9), a heavy chord along its lower edge, a

smaller chord along its upper edge, and various stiffeners. The keel beam upper chord is

attached to the WCS lower skin panel by a series of aluminum bolts/rivets forward of the

mid spar and by titanium bolts aft of the mid spar. Additionally, the keel beam upper chord

is attached to each WCS lateral beam (front spar, SWB3, SWB2, mid spar, SWB1, and

rear spar) by steel tension bolts. 

1.6.1.2 747-100 Air Conditioning Equipment Description

The 747-100 air conditioning system comprises three air conditioning packs,

manufactured by Hamilton Standard, that reduce the temperature and pressure of hot bleed

air from one or more of the airplaneís engines, the APU, or the high-pressure ground

power carts during ground operations33 to provide environmental control (pressurization,

ventilation, and temperature) to the cockpit and the main cabin. (Figure 6 is a schematic

diagram of the 747-100 air conditioning system.) The air conditioning packs are located in

an enclosed area (the air conditioning pack bay) under the WCS in the following

locations: pack No. 1 is located to the left of the keel beam (beneath the forward left

portion of the CWT); pack No. 2 is located immediately aft of pack No. 1; and pack No. 3

31 The compartment between SWB3 and the front spar was originally equipped with a bladder cell for

water and plumbing to inject that water into the engines to increase engine thrust on takeoff. When the

original engines were replaced with higher thrust engines, water for water injection was not needed, and the

compartment between SWB3 and the front spar was converted to a dry bay. The accident airplaneís water

system had been deactivated, and the bladder cell and plumbing had been removed. In some later models of

the 747 and in the military (E-4B) version, this dry bay has been modified to carry fuel. 

32 An electrical conduit is a rigid tube that contains electrical wires. 

33 According to postaccident interviews with TWA ground operations personnel, while the accident

airplane was on the ground at JFK, two of the three air conditioning packs were operated (powered by the

APU) in accordance with TWA procedures. 
Factual Information 17 Aircraft Accident Report

is located adjacent to pack No. 1 (beneath the forward right portion of the CWT) on the

right side of the keel beam (see figures 7 and 8). 

Each air conditioning pack includes the following major components:

ï A pack control valve, which controls the volume of air supplied to its

respective pack (and bleed trim system).34

ï Dual heat exchangersóall air going through the air conditioning pack must

pass through both the primary and secondary heat exchangers. Heat exchanger

cooling air is provided by ram air in flight and the air cycle machine (ACM)

fan on the ground. When minimum cooling is required, ram air passing

through the primary and secondary heat exchangers provides sufficient cooling

and the ACM is bypassed. 

ï An ACM, which consists of a fan, compressor, expansion turbine, and bypass

valve. When partially cooled air from the heat exchangers requires additional

cooling, the bypass valve closes and the air is routed through the ACM

expansion turbine for maximum cooling. 

The air conditioning packs remove heat from the engine bleed air through the

primary and secondary heat exchangers (with excess heat exhausted into the air

conditioning pack bay through a set of louvers that are flush with the lower fairings). After

conditioned air leaves each of the air conditioning packs,35 it is routed through ductwork

to the aft side of the rear spar, then along the aft side of the rear spar (through the main

landing gear wheel well) upward until it reaches the top of the CWT. The conditioned air

is then routed forward between the upper skin of the WCS and the main cabin floor into a

common plenum located above the CWT.36 The conditioned air then branches off to

vertical risers in the airplaneís side walls and ascends to the air distribution/exchange

system for the pressurized portion of the fuselage above the main cabin ceiling. (Figure 9

shows the air conditioning ducts around the CWT.) Temperature and pressure sensors

located throughout the system relay information to gauges in the cockpit, permitting the

flight crew to monitor the performance and operation of the air conditioning packs. 

34 Some of the high-temperature engine bleed air from the air conditioning pack inlet is diverted to the

air conditioning ducts on top of the WCS upper skin, where it is mixed with the conditioned air on demand

to increase the temperature of the air provided to the zone being adjusted. This diverted air is termed ìbleed

trimî air. 

35 According to the Boeing 747 Operations and Maintenance Manuals, the air conditioning pack outlet

temperatures typically range from between 35° to 160° Fahrenheit (F). 

36 The 747-100 cockpit/cabin has four air conditioning zones for which the flight and cabin

crewmembers can select independent temperatures. The three air conditioning packs then deliver air to the

plenum that has been cooled to the temperature required by the zone with the greatest cooling demand;

temperatures within the other zones can be adjusted independently by adding engine bleed trim air.
Figure 6. A schematic diagram of the 747-100ís air conditioning system. PACK2
Figure 7. Top and side views of the 747-100ís wing center section and air conditioning system.
Figure 8. Underside view of the 747-100ís wing center section and air conditioning system. Keel Beam Ducts from Packs to Plenum Pack 1 Pack 3 Bleed Air From Right Wing EnginesPack 2

Forward Spar

Spanwise Beam 3

Spanwise Beam 2

Mid Spar

Spanwise Beam 1

Rear Spar

Forward

Left Side of CWT & Dry Bay

Port for Ground

Air Source

Bleed Air

From Left

Wing Engines

Louvers
Factual Information 21 Aircraft Accident Report

Figure 9. The air conditioning ducts around the 747-100ís center wing fuel tank.

1.6.1.3 747-100 Electrical and Wiring Information

1.6.1.3.1 747-100 Electrical Information

The accident airplaneís electrical system consists of four engine-driven generators

(one per engine) and one APU-driven generator,37 all of which are capable of providing

115/200-volt,38 three-phase, 400-Hz alternating current (a.c.). These generators are driven

by a constant-speed drive located on the engine/APU gearbox, as applicable. Two

36-ampere (amp)39-hour nickel cadmium batteries provide 24-volt backup direct current

(d.c.) power and APU start power and electrical bus40 and switching logic for power

37 A second APU-driven generator is optional and is installed in some 747-100 airplanes.

38 A volt is the basic unit of measurement of electromagnetic force (the force that causes electrons to

flow through a conductor of specified resistance). 

39 An amp is the basic unit of measurement of electric current flow.

40 An electrical bus is a power distribution point to which a number of circuits may be connected. It

often consists of a solid metal strip in which a number of terminals are installed; however, it can also consist

of a section of wire from which power is distributed to other wires/circuits. 
Factual Information 22 Aircraft Accident Report

distribution. An inverter41 provides power from the batteries to flight critical equipment

when primary a.c. power is not available. The airplane also has receptacles for external

a.c. power. The APU generator control unit functions as an a.c. external power monitor

and prevents abnormal power from an external source from entering the airplane. 

During normal operation, the four engine-driven generators supply power to four

main a.c. buses. The generators are synchronized and connected together by the closing

bus tie42 and split bus breakers.43 The galley power distribution system obtains primary

115/200-volt a.c. power from the four engine-driven generators. Step-down transformers44

are used to convert some of the primary 115/200-volt a.c. power to 28-volt a.c. power,

which is then distributed to various instruments and most nongalley airplane lighting

systems by four main a.c. load buses. Five separate 75-amp transformer-rectifier units

(TRU), which are connected to each other through isolation relays, convert the a.c. power

to 28-volt d.c. power, which is used for control circuits and various d.c. components

throughout the airplane. 

1.6.1.3.2 747 Wiring InformationóGeneral and Specific to Accident Airplane

Boeingís general design and performance requirements for wires installed in the

747 (as outlined in its Drawing 60B40037) states the following: 

all parts specified in this drawing shall provide 30,000 hours of continuous

trouble-free operation when exposed to...ambient temperature extremes ranging

from a low of -65° to 250° F and combined with: 1) altitudes ranging from 1,000

feet below sea level to 50,000 feet above sea level, 2) vibration as defined in

D6-13014 ëModel 747 Equipment Vibration Test Requirements; Area 7,

Category A,í 3) relative humidity varying from 0 to 100 percent, and 4) exterior

contamination in the form of moisture, dust, sand, and metallic particles.

Boeingís definition of moisture includes salt water; Skydrol 500 A hydraulic fluid,

as described in Boeing Document D6-1100; hydraulic fluid conforming to MIL-H-5606;45

aviation lubricating oil conforming to MIL-L-7808; and jet fuels conforming to

MIL-J-5624, D1655-59, P&WA 522, or caustic cleaning fluid.

Early production 747s were manufactured using Boeing Material Specifications

(BMS)13-38 or BMS13-39 wiring for all nonfuel tank applications; however, Boeing

41 According to the FAA Airframe & Powerplant Mechanics General Handbook, Advisory Circular

(AC) 65-9, an inverter is a mechanical or electronic device used to convert d.c. power to a.c. power, which is

then used ìmainly for instruments, radio, radar, lighting, and other accessories.î 

42 A closing bus tie is used to electrically connect two or more electrical buses. 

43 The split bus breaker is a manually operated circuit breaker used to separate the a.c. power

distribution system (which normally operates with the four a.c. buses joined together as one electrical power

bus) into two separate a.c. buses. 

44 A step-down transformer is used to reduce higher-voltage a.c. power for use in airplane systems that

require lower voltages. 

45 The designation ìMILî is used to identify components and parts manufactured to the standards

described in Department of Defense (DoD) documents known as Military Standards. MIL parts are

commonly used in civilian transport airplanes. 
Factual Information 23 Aircraft Accident Report

adopted the use of BMS13-42 wiring on July 30, 1969. BMS13-42 wiring is an aliphatic

polyimide (also termed ìalkane-imideî) insulated wire, also known by the trade name

Poly-X,46 and was manufactured by Raychem.47 Boeingís Manufacturing Development

Report 6-27037, dated April 5, 1970, stated, ìEmployment of the new wire was

designatedÖbecause its thinner insulation provided a potential weight saving of

approximately 400 to 600 pounds per airplane over that achieved with the existing

BMS13-38 and BMS13-39 wire.î 

On April 13, 1970, Boeingís 747 material specification for Poly-X wire of the

BMS13-42/8 type used in nonfuel quantity indication system wire bundles was revised

slightly to specify the use of BMS13-42A.48 BMS13-42A Poly-X wire consists of three

layers of modified alkane-imide polymer coating (0.015 inch minimum thickness) over a

tin-coated copper core conductor and inner and outer layers of primary insulation

(0.009 inch minimum thickness). Boeingís material specification for this type of wire

states that ìa coating of modified imide polymer shall be applied over the insulation. This

coating shall be continuous and free from cracks, splits, blisters, and other defects when

examined without the aid of magnification.î49 Figure 10 shows a BMS13-42A Poly-X

wire, with typical markings specified by Boeing. 

A wire type table contained in the FAAís Aging Non-Structural Systems Program

documentation indicated that the use of Poly-X wire was discontinued in 747 production

in 1975.50 According to FAA records, Boeing used MIL-W-81044/20 (also identified by

Boeing as BMS13-42C and -42D and known by the trade name Stilan) in 747 production

between 1975 and 1979. The records indicate that Boeing has used cross-linked

ethylenetetrafluoroethylene (XL-ETFE, per MIL-W-22759; also identified by Boeing as

BMS13-48 and known by the trade name Tefzel)51 in 747 production in various

applications since 1979. However, the records also show that Boeing used MIL-W-81381

(also identified by Boeing as BMS13-51 and known by the trade name Kapton)52 in 747

production between 1985 and 1993. Most of the wiring in the accident airplane was

BMS13-42A Poly-X wiring and was stamped with a green ì42A,î although other types of

46 The term ìPoly-Xî also applies to BMS13-42A, an aliphatic polyimide insulated wire with a thicker

external coating (on some larger diameter wires) than BMS13-42. For simplicity, in this report, the term

ìPoly-Xî will be used to indicate BMS13-42/42A wiring in general. Where it is appropriate, the report will

refer specifically to either BMS13-42 or BMS13-42A Poly-X. 

47 Although Raychem stopped manufacturing Poly-X wire in 1975, the current Boeing Standard Wiring

Practices Manual (SWPM), 20-00-13, table III, page 96, continues to specify Poly-X as an approved

alternative material for use. 

48 BMS13-42 was still used in woven sleeves and other protected applications. 

49 Boeing required specialized tools for use with Poly-X wiring and warned that use of other tools ìwill

result inÖdamage to conductor or insulation.î For additional information regarding the condition of the

wiring recovered from the accident airplane and wiring in other inspected airplanes, see section 1.18.2. 

50 According to the FAAís Aging Non-Structural Systems Program documentation, ìairplane

manufacturers typically continue to use existing stock until it is exhausted, [and as a result, some material]

changeovers may have taken considerable time. Thus, trying to determine the wire type installed based on

the date of manufacture of an airplane is not necessarily accurate.î 

51 For simplicity, in this report, XL-ETFE will be used to indicate cross-linked ETFE, per

MIL-W-22759. 
Factual Information 24 Aircraft Accident Report

wire were specified and recovered. For example, the fuel quantity indication system

(FQIS) wiring used in the fuel tanks was polytetrafluoroethylene (PTFE), per

MIL-W-16878, which is also known by the trade name Teflon.53

52 For simplicity, in this report, the trade name Kapton (a du Pont trade name for a material that has

multiple applications) will be used to indicate MIL-W-81381 wire. During an April 1999 presentation,

du Pont showed that this type of wire had good abrasion/cut-through (resistance), low-smoke/nonflame, and

weight and space characteristics. Du Pont indicated that its limitations were arc-track resistance and

flexibility. 

53 According to Boeingís 747 specifications, PTFE (or Teflon) has been used for FQIS wiring in fuel

tanks and between the CWT and gauges at the flight engineer station in the cockpit. For simplicity, in this

report, the trade name Teflon will be used to indicate PTFE wiring, per MIL-W-16878. 

W103

-RN

-1234

20

denotes Wire Bundle Number

denotes System Function Number

denotes Wire Serial Number

denotes Gauge

BMS13-42A-type

Wire with Double Insulation Layer and Topcoat

Figure 10. Diagram of a BMS13-42A Poly-X wire.
Factual Information 25 Aircraft Accident Report

1.6.1.3.3 Separation of 747 Wire Circuits

The Safety Boardís review of FAA regulations revealed that the FARs do not

contain specific guidance regarding separation of electrical circuits/wiring. However,

14 CFR Section 25.1309 states that ì[t]he airplane systems and associated components,

considered separately and in relation to other systems, must be designed so thatÖthe

occurrence of any failure condition which would prevent the continued safe flight and

landing of the airplane is extremely improbable.î54

 Boeing Document D6-24391, ìCriteria for Separation of Critical Electric

Circuits,î dated March 17, 1970, provided guidelines for separation that Boeing deemed

necessary for isolation of critical systems. These guidelines were applicable to ìany

electrical equipment or system for which the proper functioning is considered essential to

safe operationî in all Boeing commercial airplanes at that time, including the 747-100

airplane. The document stated that the ìobject of the circuit separation is to prevent

hazardous malfunctions or simultaneous loss of redundant power supplies or redundant

equipment functions due to failures such as: 1) fire or damage to any wire bundle, 2) loss

of any single connector, 3) fire in a junction box, or 4) engine turbine burst.î 

D6-24391 listed the circuits to separate in each major airplane system, the reasons

for the separation, and the methods to be used for the separation. For those systems

designated as protected (such as fire warning and protection system wiring, wiring to

individual fuel pumps, and individual engine installation circuits), Boeing specified

separation distances of at least 1/4 inch in pressurized areas and at least 1/2 inch in

unpressurized areas.55 According to Boeingís SWPM, where such separation is not

possible because of space or other constraints, separation by an insulation material (as

described in the SWPM) or by a fusible link circuit breaker is required. 

Although D6-24391 and Boeingís SWPM indicated that wiring to individual fuel

pumps needed to be protected by separation, no special separation was required for other

FQIS and fuel system wiring, including the fuel quantity probes and interconnecting

wiring and the fuel quantity gauges. However, Boeing incorporated an electromagnetic

interference (EMI) shield on the FQIS wire bundle between the flight engineerís panel and

the CWT on 747s produced after line number 243 (the accident airplane was line

number 164).56

54 For more information about 14 CFR Section 25.1309, see section 1.16.6.5. 

55 In a May 25, 1999, letter to the Safety Board, Boeing indicated that it had conducted a series of tests

during the design and certification of the 747 in 1969 and 1970, and again in 1980, that indicated that wire

bundles separated by a minimum 1/4-inch air gap from a failed bundle will not sustain damage that

compromises the electrical integrity of the wire bundle. 

56 According to a May 26, 1998, letter to the FAA, Boeing stated that the shielding was added ìdue to

fluctuating indication caused by EMI.î EMI is further discussed in section 1.16.6.4 of this report. 
Factual Information 26 Aircraft Accident Report

The Safety Boardís review of other manufacturersí separation standards revealed

the following: 

ï In Douglas Aircraft Company57 specification WZZ-7002, Douglas specified

that wiring for protected systems (including FQIS and other fuel system

wiring, fire warning system wiring, primary generator feeder cables, and

electro-explosive devices) must be isolated by at least 3 inches from other

electrical wiring. Further, Douglas specified that the routing of such wires must

be continuous from the tank to the cockpit gauge, with no connectors58 shared

with other systems. 

ï In a November 2, 1998, letter to the Safety Board, Airbus stated that ìthe

design aim for AIRBUS aircraft has been for all in-tank wiring to be

segregated from power sources. As such, electrical segregation exists on all

aircraft together with physical segregation from 115-volt power wiring. We

have found an exception with some fuel [quantity] sense wiring in A300 and

A310 [airplanes] where wiring runs locally in the same routes as 115-volt

wiring. This point is under reviewÖî

1.6.2 747-100 Fuel System Description

The 747-100ís fuel system consists of seven fuel tanks (three in each wing and one

in the CWT), an engine crossfeed system (manifold and valves),59 a fueling station on the

left wing, a fuel jettison system, and a surge tank in each wing tip. The fuel tanks include

the inboard and outboard main tanks and a reserve tank in each wing and the CWT (which

occupies most of the WCS in the fuselage). According to Boeing, the wing fuel tank

capacities for the 747-100 are as follows: reserve tanksó3,350 pounds (500 gallons)

each; outboard main tanksó29,614 pounds (4,420 gallons) each; inboard main tanksó

82,008 pounds (12,240 gallons) each; and the CWTó86,363 pounds (12,890 gallons).

Figure 11 is a schematic diagram of the 747-100 fuel tank arrangement.

The left and right wing fuel tanks are formed by sealing the internal wing structure

so that the front and rear spars define the forward and aft limits of the tanks, respectively.

The left and right wings also contain dry bays, which are located outboard of the reserve

wing tanks. As previously discussed, the forward and aft limits of the CWT are defined by

SWB3 and the rear spar, respectively; the space between SWB3 and the front spar in the

WCS is a dry bay. A sump drain is located near the lowest point of each of the seven fuel

tanks.

57 Douglas Aircraft Company merged with Boeing and is now the Douglas Products Division of the

Boeing Commercial Airplane Group. 

58 A connector is a device that permits in-line connection(s) between one or more wires for a continuous

electrical path(s) at a location where the wires may be disconnected and reconnected without mismatching

circuits. Typical multicontact electrical connectors are assembled from two subassemblies, the plug and the

receptacle, which mate to connect wires with pin and socket contacts. 

59 The fuel crossfeed system permits the transfer of fuel from any fuel tank to any engine; however, it

does not permit the transfer of fuel from one fuel tank to another fuel tank. 
Figure 11. A schematic diagram of the 747-100 fuel tank arrangement.
Factual Information 28 Aircraft Accident Report

Each wing also contains a surge tank, located outboard of the dry bays. The surge

tanks in the wing tips act as reservoirs (vent collector tanks) if the other fuel tanks are

overfilled. The top of each fuel tank is vented to outside atmosphere. The vent system

consists of continuous hollow spaces inside box-shaped stringers attached to the interior side

of the upper skins of the wing and CWT. These vent stringers are routed outboard from the

top of each fuel tank in the wings to the surge tank tip of that wing. The CWT is vented

directly to the surge tanks in both wings. To complete the venting system path to outside

atmosphere, a single tube connects the outboard side of each surge tank with the outside

atmosphere at the lower wing tip surface.60 Figure 12 shows a typical 747-100 CWT and left

wing fuel tank vent system layout. The right wing fuel tank vent system layout is similar.

To protect against the propagation of a flame from the surge tank to any of the

other fuel tanks through the vent system,61 the 747-100 surge tank overboard vents are

equipped with a surge tank protection (STP) system. The STP system consists of an

optical photocell located in each overboard vent designed to trigger the discharge of Halon

(a fire extinguishing agent) into the surge tank when a flame or bright light source62 is

sensed in that overboard vent. Discharge of the extinguishing agent is designed to occur

about 1 millisecond after the photocell senses a flame.

The 747-100 uses a pressure fueling system to distribute fuel under pressure from

a fueling station in the left wing63 to all of the fuel tanks through a crossover manifold, the

main distribution manifold,64 and refuel valves (one each at the reserve and outboard fuel

tanks and two each at the inboard fuel tanks and the CWT). The fueling station is equipped

with two fueling receptacles coupled together and connected to the crossover manifold,

which extends through the outboard fuel tank to the main distribution manifold. In

addition, a fueling control panel is located in the lower leading edge of the left wing, near

the fueling station where ground service hoses connect to the refueling ports. Figure 13

shows the 747 pressure fueling flow diagram, excerpted from Boeingís 747 Maintenance

Manual.

60 According to Boeing, the vents are sized to permit passage of fuel if an overfill situation occurs

during ground refueling. This design criterion requires a larger vent cross-section area than would be

required to provide air pressure equalization alone. 

61 Flames entering fuel tanks through the vent system was a factor in the December 8, 1963, accident

following a lightning strike on a Pan Am Boeing 707 in Elkton, Maryland, and in the November 23, 1964,

accident involving a TWA 707 in Rome, Italy. 

62 According to a representative from Fenwal Safety Systems (the manufacturer of the STP system), the

system has been triggered by lightning, flamefronts, and even bright flashlights (used by mechanics during

maintenance or by pilots during preflight inspections). 

63 Although the accident airplane only had a fueling station on the left wing, at the time of its

manufacture, an optional right wing fueling station was available and installed on some 747s.

64 The main distribution manifold, which extends through the length of the wing, terminating at a fixed

jettison nozzle on the wing trailing edge near each wing tip, is also used as the fuel jettison manifold. Fuel

can only be jettisoned from the inboard main tanks and the CWT.
Figure 12. A layout of a typical 747-100 center wing fuel tank and left wing fuel tank vent system.
Factual Information 30 Aircraft Accident Report

CENTER WING TANK

DRY BAY

M

M

M M

P R

R

R

2

1

R

R

R

R

OVERRIDE/JETTISON

PUMP PRIMING LINE

REFUEL

VALVE

REFUEL

VALVE

JETTISON

NOZZLE VALVE

(REF)

SURGE

TANK

RESERVE

TANK

NO. 1

MAIN

TANK

NO. 1

MANIFOLD DRAIN

VALVES (2 PLACES)

PLUGS

(2 PLACES)

REFUEL

VALVE

REFUEL

VALVE

REFUEL

VALVE

CROSSOVER

MANIFOLD

MAIN DISTRIBUTION

MANIFOLD

MANIFOLD

DRAIN VALVE

MANUAL

SHUTOFF VALVE

FUELING

RECEPTACLES

MANUAL

DEFUELING VALVE

MAIN

TANK

NO.2

CROSSFEED MANIFOLD

CHECK VALVE

REFUEL

VALVE

PRESSURE FUELING

ALTERNATIVE INSTALLATION

PREFERRED INSTALLATION

REFUEL

VALVE

OVERRIDE/JETTISON

PUMP (REF)

JETTISON

VALVE

(REF)

M

P

Figure 13. A 747 pressure fueling flow diagram.
Factual Information 31 Aircraft Accident Report

The wing refueling control panel contains fuel quantity indicators,65 refuel valve

control switches, valve position indicator lights, a refuel power switch, and a test switch

and indicators for the volumetric shutoff (VSO) control. (Figure 14 shows the pressure

fueling system components.) The VSO control unit can be set to automatically close the

refuel valves when a preselected amount of fuel is loaded. The preselected amount of fuel

can range in the tanks from any partial quantity to full fuel tanks.66 Refueling can also be

controlled by ground personnel, who can close the refueling valve manually when the fuel

quantity indicator shows that the desired quantity of fuel is on board the airplane. The

VSO will also shut the valves when fuel enters the surge tanks if a fuel tank is overfilled.

1.6.2.1 Fuel Quantity Indication System Components and Wiring 

Information

The amount of fuel in each of the 747-100ís seven fuel tanks is measured by a set

of FQIS probes and compensators,67 with the number of fuel quantity probes and

compensators differing for each tank. The 747-100 has 65 fuel quantity probes and 13 fuel

quantity compensators. The fuel quantity probes are positioned as follows: 4 each in the

Nos. 1 and 4 reserve fuel tanks; 13 each in the Nos. 1 and 4 main fuel tanks; 12 each in the

Nos. 2 and 3 main fuel tanks; and 7 in the CWT. The fuel quantity compensators, located

near the low point of each fuel tank, are positioned as follows: two in the No. 1 reserve

tank; one in the No. 4 reserve tank; one in each surge tank; two each in the Nos. 1 and

4 main fuel tanks; two in the No. 2 main fuel tank; one in the No. 3 main fuel tank; and

65 The fuel quantity indicators at the wing refueling control panel are repeater type indicators; they share

information with corresponding primary indicators on the flight engineerís lower instrument panel for

display in the cockpit. For additional information regarding FQIS components/indicators, see

section 1.6.2.1.

66 While the accident airplane was being fueled at JFK, the fuel systemís automatic VSO activated

before the fuel tanks were full. According to postaccident interviews, a TWA mechanic examined the fuel

system and ì[overrode] the system, pull[ed the] volumetric [shutoff valve or VSO] fuse and an overflow

circuit breaker.î After the fueler finished fueling the airplane manually, the TWA mechanic reset the fuse

and circuit breaker. According to the fueler and TWA maintenance personnel, activation of the VSO is a

common occurrence. According to maintenance records, the accident airplane had numerous VSO-related

maintenance writeups during the weeks before the accident. For additional information about the accident

airplaneís previous refueling difficulties, see section 1.6.3. 

67 The 747-100 FQIS measures changes in the capacitance of tubular fuel quantity probes located in

each fuel tank for the display of fuel quantity on cockpit indicators and on repeater indicators located at the

refueling station in the left wing. The system is also connected to systems that require fuel quantity

information, including the gross weight/total fuel weight indicator, airborne integrated data system (AIDS),

and VSO (ground refueling) systems. Fuel quantity probes are components of the FQIS, a set of which is

positioned in each of an airplaneís fuel tanks to measure the quantity of fuel in each tank through a range of

levels within the tank. Fuel quantity compensators are used in the FQIS and VSO systems. For FQIS

applications, compensators are used to compensate for variations in the electrical properties (the dielectric

constant) of fuel, which vary from one type of fuel to another (and even within the same type of fuel,

depending on the ìbatchî and the age of the fuel) to ensure consistent fuel quantity indications. About 90

percent of the 747-100, -200, and -300 series airplanes are equipped with FQIS components manufactured

by Honeywell Corporation; the remaining 10 percent of the 747-100, -200, and -300 series airplanes and all

747-400 airplanes have been retrofitted with FQIS components manufactured by BFGoodrich Aerospace

Corporation (formerly Simmonds Precision) or Smiths Industries. 
Figure 14. A diagram of the 747 pressure fueling system components.
Factual Information 33 Aircraft Accident Report

one in the CWT. Figure 15 shows a fuel quantity probe and a compensator. Figure 16

shows the fuel quantity probe and compensator locations in the CWT.

The fuel quantity probes in the 747-100 fuel tank contain two metallic electrodes,

which are assembled as inner and outer tubes and are oriented vertically within the probe.

A hard plastic terminal block is located near the top of the fuel quantity probe for wiring

connections, and two brackets mount the probes vertically in the fuel tank. The diameter

of the inner electrode varies, based on the contours of the fuel tank, and the length of each

fuel quantity probe is determined by the height of the fuel tank at that location. Because

the capacitance68 of air is different than that of fuel, capacitance values for each fuel

quantity probe are different when air, instead of fuel, fills the gap between the inner and

outer tubes. Thus, the capacitance of the fuel quantity probe varies linearly, depending on

the depth of fuel in the tank.

The fuel quantity compensators in the 747-100 fuel tank contain three metallic

electrodes, an external mounting tube, a hard plastic terminal block near the top of the

mounting tube for wiring connections, and two mounting clamps. A fuelís density (and,

therefore, its dielectric constant) changes with temperature and other variables; the

compensators provide a ì0î reference for the FQIS. Therefore, the fuel quantity

compensators are mounted at the lowest point in the fuel tank so that they remain in

contact with fuel until the tank is almost empty.

The fuel quantity probes and compensators in the 747-100 fuel tanks are connected

in series by wires, and the entire set is then connected to a fuel quantity gauge on the flight

engineerís panel in the cockpit. The cockpit fuel quantity gauge generates an electrical

voltage, and the resulting electrical capacitance between the probesí inner and outer tubes

is measured. The measurements of fuel in the gaps between the tubes are summed, and the

amount is calibrated so that the flight engineerís fuel quantity gauge indicates the weight

of the remaining fuel instead of the capacitance. (Weight is the common measurement for

fuel quantity in transport airplanes.) A gauge at the left wing pressure refueling control

panel repeats at the cockpit display.

Although most of the wiring for the accident airplane was BMS13-42A Poly-X,

the wiring used for the 747 FQIS wires in the fuel tanks and from the CWT to a wire splice

located behind the flight engineerís panel was silver-plated copper wiring with Teflon

insulation. Specifically, Teflon wires were routed from the fuel quantity probes and

compensators in the fuel tanks, through the interior of the fuel tank supported by

numerous nylon clips, to the electrical connector located on the aft side of the rear wing

spar for the CWT and on the forward side of the forward wing spar for the main fuel tanks.

Outside of the fuel tanks, FQIS wires were routed from the connectors on the wing spars

to a splice located near each of the fuel quantity indicators (one for each fuel tank) located

on the flight engineerís panel in the cockpit,69 then to the fuel quantity totalizer indicator

and the VSO unit.70 Seven fuel quantity repeater gauges are mounted at the left wing

refueling panel;71 wires to these gauges are routed through the wing fuel tanks from the

68 Capacitance is the property of conductors separated by a dielectric material (for example air or fuel)

that permits the storage of electricity when potential differences exist between conductors. 
Factual Information 34 Aircraft Accident Report

cockpit gauges. The only electrical wiring/components located inside the 747 CWT (or

any 747 fuel tanks) are those associated with the FQIS (including the ground refueling

system).72 According to representatives from Honeywell and Boeing, electrical power is

supplied to the FQIS components through the cockpit fuel gauges and is intended to be

limited to 0.02 millijoule (mJ).73

Wiring diagrams for the 747 indicate that FQIS and fuel flow indication wire

bundles74 were routed along common raceways75 from the flight engineer station to the

electrical equipment compartment located between the nose landing gear wheel well and

the forward cargo compartment. The diagrams also show that the left wing and CWT

FQIS wires were routed in a raceway with a wire bundle containing wires leading to the

CVR for part of their path through the upper left fuselage and adjacent to a wire bundle

containing wires leading to the right main wing fuel tank for part of their path behind the

flight engineer station (upper right fuselage). Additionally, the diagrams indicate that, if

Boeingís PI 61B7010376 were followed for the accident airplane,77 the left wing FQIS

wires would have been corouted in a raceway with wires from other systems (such as

lights) that were bundled with both the CVR and the No. 4 fuel flow indicator wires for a

portion of their paths. Further, the wiring diagrams show that the right wing FQIS wires

were routed to a connection at the T347 terminal strip in the CWT. Figure 17 is a

69 The primary indicators convert fuel quantity probe and compensator capacitance into dial and counter

indications of fuel quantity. Each primary indicator has two variable resistor adjustment shafts for

calibrating the indicator so that it displays the proper zero and full indications for empty and full fuel

quantity probe values. 

70 The fuel quantity totalizer indicator (termed the ìgross weight/total fuel weight indicatorî in the TWA

747 Operations Manual) is a dual-counter display device located on the flight engineerís panel. It converts

signals from the seven primary fuel quantity indicators into an indication of total fuel weight, which is

displayed on the lower of the two counters. The second counter display indicates the airplaneís gross weight. 

71 Each repeater indicator converts a signal from the associated primary indicator into a duplicate

indication of fuel quantity. The repeater indicators are used for refueling operations and are deenergized

whenever the door to the refueling panel is closed. Each repeater indicator has two variable resistor

adjustment shafts for calibrating the indicator so that it displays the proper zero and full indications for

empty and full fuel quantity probe values. 

72 Although the fuel pump motor cavities contain wiring, they are separated from the CWT by flame

suppression passages and other parts. 

73 A joule (J) is a unit of measurement of electrical work or energy; 1 J is the amount of work done by

1 watt of power in 1 second. The power supplied to the FQIS components through the cockpit fuel gauge

(0.02 mJ) is about 10 percent of the minimum ignition energy (MIE) requirement (0.25 mJ) for hydrocarbon

fuels, referenced by the American Petroleum Institute Recommended Practice 2003 (API 2003), ìProtection

Against Ignitions Arising out of Static, Lightning, and Stray Currents,î fifth edition, December 1991. 

74 The Boeing SWPM refers to wires that are tied in a group as a ìharnessî; the TWA Master Wire

Bundle Document and Boeing production illustrations (PI) refer to wires that are tied in a group as a

ìbundle.î In this report, the term ìbundleî will be used to refer to wires that are tied in a group, except where

a reference source is being quoted. 

75 ìRacewayî is the term used for areas in the 747 where wire bundles are grouped into a common route. 

76 The most recent revision to PI 61B70103 was dated November 10, 1971. 

77 According to an October 13, 1999, letter from Boeing, the original wiring installation diagram had

been superseded by a more direct wiring path for the No. 4 fuel flow indicator wires when the accident

airplane was manufactured. No documentation of this change was provided. The actual wiring route of the

No. 4 fuel flow indicator wires in the accident airplane could not be determined. 
Factual Information 35 Aircraft Accident Report

Figure 15. Diagrams of a fuel quntity probe and a compensator.
FQCO

Figure 16. The fuel quantity probe and compensator locations in the 747ís center wing fuel tank.
Factual Information 37 Aircraft Accident Report

schematic diagram of the FQIS wiring, fuel flow wiring, CVR wiring, and other wiring

information from the 747-100, and figure 18 shows the locations of the wire raceways in

the 747-100. 

Boeingís PIs also indicated that the CWT FQIS wires were routed from the front of

the airplane to the CWT, VSO, totalizer, and the AIDS in and next to bundles carrying

numerous other power and signal wires. Although most airplane systems operated on

115-volt power, the Safety Boardís review of wiring in other airplanes78 indicated that the

FQIS and engine fuel flow wire bundles are routed near, and occasionally bundled with,

wires that provide 350-volt a.c. power to the cabin fluorescent lights.79

According to TWA, the fuel quantity probes and compensators (and most other

electrical and fuel system components, including wiring and connectors) are

condition-monitored items, which means that they are removed and replaced only when

inoperative. Unless a malfunction of the FQIS requires unscheduled maintenance, the

units are inspected only during the scheduled D check.80 According to TWA maintenance

records, the fuel quantity probes and compensators installed in the accident airplaneís fuel

tanks were the original units installed when the airplane was manufactured in 1971, except

for four of the fuel quantity probes in fuel tank No. 1, which records indicate were

replaced during a 5B check in September 1987.81 

1.6.2.2 747-100 Fuel Pump System Description 

The 747-100 has 16 fuel pumps, including 4 fuel boost pumps in each wing (2 per

engine),82 2 fuel jettison/override pumps in each wing,83 2 fuel jettison/override pumps in

78 For more information about the Safety Boardís review of wiring in other airplanes, see section 1.18.2.

79 Although the Safety Board is not aware of instances in which FQIS wires shared connectors with

wires carrying more than 115 volts, Boeing specifications permit FQIS wiring to be mixed in common

connectors with airplane system wires carrying up to 350 volts. 

80 For additional information regarding the scheduled D check, see section 1.6.3.1.

81 Although there was no maintenance record indicating that any other fuel quantity probes or

compensators had been replaced, an April 20, 1993, maintenance record indicated that maintenance was

conducted to correct fuel quantity discrepancies in the No. 1 main fuel tank; this maintenance record showed

capacitance measurements for each No. 1 main fuel tank probe. There was no record of probe removal with

this maintenance action. According to the Maintenance Records Group report, a review of maintenance

records from other TWA airplanes indicated that malfunctioning probes were removed and sent to the

overhaul shop for cleaning. 

82 Fuel boost pumps move fuel from the fuel tanks to the engines or the APU. 

83 The jettison/override pumps have two uses in the 747, one of which is to jettison fuel from the

airplane when this function is selected at the flight engineer station. The jettison/override pumps also

provide fuel to the engine manifolds at a higher pressure than the fuel boost pumps; thus, fuel pressure from

the jettison/override pumps in the CWT closes the main fuel tank check valves, ensuring that the fuel from

the CWT is consumed before the fuel in the main wing fuel tanks. 
Fragments ofVSO and FF4wires(Ref. Sta 360frame)

Main deck floor (Ref. WL200)

Flight EngineerStation (Ref.)

Volumetric Shutoff Control Unit (VSO)(Ref. E3-1 Rack, Sta. 460)

Fuel Flow Computer(FF4) (Ref. E3-1 Rack)

Cockpit floor edge(Ref.)

Fuel totalizer

Fuel quantity gauge

Right wing fuel 

tank wire route

(ref.)

Wire routes 

wing fue

refuelin

Tank Unit (probe)

Fragments 

16

15

13

13

12

1

Aircraft integrated

data system (AIDS)

(Ref. E11 Rack, Sta. 472)

Background illustration shows mabundle raceways & identification 

Right wing FQIS & fuel 

flow #4 wire routes (ref.)

CVR wire route (ref.)

FF#4 wire route shown 

on Production Illustration (ref.)

Compensa

Figure 17. A schematic diagram of the 747-100ís fuel quantity indication system wiring, fuel flow wiring, cockpit vwiring, and other wiring information. 
FLIGHT ENGINEER STATION H IK SHOWN IJ OPPOSITEIM SHOWN IN OPPOSITECC P G CWT RE

SPAR CFOR FQ

Figure 18. Locations of the wire raceways in the 747-100. 
Factual Information 40 Aircraft Accident Report

the CWT, 1 fuel scavenge pump in the CWT,84 and an APU d.c. fuel pump in the left wing.

(Figure 19 is a schematic diagram of the fuel pump locations in the 747.) The fuel

scavenge pump, APU fuel pump, and fuel boost pumps are condition-monitored items.

According to TWA maintenance records, TWA established a hard-time life limit of 9,000

hours for the fuel jettison/override pumps.85

According to TWA, the companyís maintenance facility in Kansas City, Missouri,

conducted all 747 fuel pump overhauls and repairs in accordance with the manufacturerís

overhaul manuals and (as in the case of the jettison/override pumpís 9,000-hour hard-time

limit) specifications established by TWAís engineers and accepted by the FAA.

Examination of the maintenance history of the fuel pumps installed in the accident

airplaneís CWT at the time of the accident revealed the following:

ï Fuel scavenge pump, CWTóS/N M285; this pumpís most recent overhaul was

signed off on October 18, 1994, and the unit passed a functional test on

October 20, 1994. The pump was installed on the accident airplane on

April 22, 1996. At the time of the accident, the pump had accumulated

1,040 flight hours since overhaul. 

ï Jettison/override pump, CWTóS/N 690469A; this pumpís most recent

overhaul was signed off on March 14, 1996, and the unit passed a functional

test on March 18, 1996. The pump was installed on the accident airplane on

March 27, 1996. At the time of the accident, the pump had accumulated

1,270 flight hours since overhaul and had an estimated (9,000-hour) removal

date of February 10, 1998. 

ï Jettison/override pump, CWTóS/N 690394; this pumpís most recent overhaul

was signed off on December 6, 1994, and the unit passed a functional test on

December 8, 1994. The pump was installed on the accident airplane on

May 15, 1995. At the time of the accident, the pump had accumulated

4,412 flight hours since overhaul and had an estimated (9,000-hour) removal

date of June 24, 1997.

84 The scavenge pump is designed to remove the last amounts of usable fuel from the CWT and pump it

into the left inboard (No. 2) fuel tank. The intake opening for this pump is located very close to the bottom of

the tank between the mid spar and SWB2. Although the scavenge pump removes fuel that is unobtainable by

the jettison/override pumps, a small amount of residual fuel will remain in the fuel tank that the scavenge

pump is not able to remove (unusable fuel). In the 747 CWT, the jettison/override and scavenge pumps are

mounted on the rear spar such that the pump motor is external to the tank but contains fuel-cooled windings.

Fuel for cooling is supplied to the windings through a small-diameter tube in series with a check valve, both

of which are intended to suppress flames. The pumps also contain a thermal fuse designed to open at less

than 400o F. 

85 More than 17 years before the accident, TWA established the hard-time limit to reduce the number of

unscheduled removals, after the company determined that the mean time between unscheduled removals for

the pumps was 11,352 hours. The 9,000-hour scheduled removal occurs at every second C check and at

every D check. For more information about C and D checks, see section 1.6.3.1. 
Factual Information 41 Aircraft Accident Report

Figure 19. A schematic diagram of the fuel pump locations in the 747.
Factual Information 42 Aircraft Accident Report

1.6.2.3 747-100 Fuel Flow Indication Information

According to TWAís Maintenance Manual, the 747-100 has a fuel flow

sensor/transmitter mounted on each of the airplaneís four engines. Fuel flowing through

each sensor/transmitter turns an impeller; information regarding the rotational speed of the

impeller is transmitted and displayed in the cockpit to indicate the fuel flow. Other major

components of the indication system are a fuel flow computer located adjacent to the VSO

computer in the electrical/electronics (E/E) bay, a fuel flow indicator located near the fuel

quantity indicators at the flight engineer station, and a repeater indicator located in the

cockpit near the center of the forward instrument panel.

The fuel flow components are connected by wires that are routed from the engines

to the main fuselage through raceways along the forward wing spars. The wires are then

routed forward through the fuselage, in raceways beneath the cabin floor, to the computer

in the E/E, then along the fuselage frame at STA 360, and upward to the flight engineerís

station. The wires routed from the fuel flow and VSO computers to the cockpit are tied

into larger bundles that include wires from other systems (including cabin lighting wires,

which carry voltage as high as 350-volts a.c.).86 Figure 17 shows the routing of fuel flow

wires through the 747-100.

In a May 2, 2000, letter, Boeing stated that the following are possible causes for an

erratic fuel flow indication (as referenced in the CVRósee section 1.1 or appendix B) in a

747-100:

1. The aircraft transmitters had low-speed motors, and the motor drive circuitry in

the signal conditioner was failing;

2. Significant contamination was present in the affected transmitter;

3. There was a fault in the signal circuitry; or

4. There was a fault in the indicator.

Further, with regard to the fuel flow meters in the 747-100, Boeingís letter stated

that ìthe portion of the mechanism which is in fuel is lubricated by fuel. The motor has

sealed bearings, which [are removed and replaced only when inoperative], not on a

schedule.î

1.6.2.4 TWA Flight 800 Fueling Information

TWAís fueling records indicated that the accident airplane was refueled at Athens

and that its fuel load for the trip to JFK included a full CWT. TWAís records further

indicated that upon arrival at JFK, the accident airplaneís CWT contained about

300 pounds of fuel; when the airplane was refueled at JFK, no additional fuel was added

to the CWT.87 According to TWAís fueling records, the accident airplane was fueled at

86 Three hundred fifty volts is the root mean square voltage (average amplitude) of the 400-Hz electric

waves in the cabin lighting circuitís wires. The peak voltage of a 350-volt a.c. circuit is slightly more than

450 volts. 
Factual Information 43 Aircraft Accident Report

JFK using the underwing pressure method. Table 2 shows TWAís fueling records, which

indicate TWA flight 800ís fuel loading when it prepared to depart JFK.

On the basis of the airplaneís dispatch fuel load information and performance data,

the Safety Board calculated that the fuel on board the airplane at takeoff was about

176,600 pounds (consistent with TWAís dispatch records) and at the time of the accident

was about 165,000 pounds.

1.6.3 Maintenance Information

1.6.3.1 TWAís 747 Maintenance Inspection Program

During its investigation of the TWA flight 800 accident, the Safety Board reviewed

TWAís maintenance program and maintenance recordkeeping procedures and conducted a

detailed examination of the accident airplaneís maintenance records. The Boardís review

of TWAís FAA-accepted continuous airworthiness maintenance inspection program for its

747 fleet revealed that it included the following inspections and intervals:

87 During postaccident interviews, the flight engineer from the Athens-JFK flight (TWA flight 881) told

Safety Board investigators that all of the usable fuel in the CWT was used during the flight to JFK and that

the fuel pumps appeared to work normally. He stated that he did not observe any change in the CWT fuel

quantity indication after the usable fuel was consumed. (According to Boeing, the CWT capacity includes

50 gallons [300 pounds] of unusable fuel.) 

Table 2. TWAís fueling records, indicating TWA flight 800ís fuel loading 

when it prepared to depart JFK.

Fuel Tank

Indicated Fuel Load Before 

Fueling (pounds)

Indicated Fuel Load After 

Fueling (pounds)

No. 1 Reserve 0 3,400

No. 1 Main 5,400 24,600

No. 2 Main 5,300 62,900

CWT 300 300a

a. The cockpit indicator has an analog scale and a digital (rotating drum) scale. When the digital 

display indicates 300 pounds, the analog needle would indicate near ì0.î 

No. 3 Main 6,900 62,700

No. 4 Main 6,300 24,600

No. 4 Reserve 0 3,300

Total 24,200 181,800
Factual Information 44 Aircraft Accident Report

ï Periodic service (PS) inspectionóincludes the cockpit and cabin, airplane

exterior, and a final check. Accomplished at scheduled layovers of 6 hours or

more at stations staffed by TWA mechanics, or every other operating day

maximum, unless a higher maintenance level is accomplished. The accident

airplaneís most recent PS was accomplished on July 17, 1996, at JFK before

the departure of TWA flight 800. 

ï Aircraft service (AS) inspectionóincludes all PS items, plus routine AS

inspection items (including checks for cockpit and cabin; exterior fuselage,

wings, gear, and empennage; engines and struts; oil, hydraulic fluid, and water

servicing; and fuel tank sumping), which are identified as AS-1 through AS-6.

AS-1 through AS-6 are accomplished in sequence at intervals not to exceed

100 hours of airplane time in service, until a time control service (TCS)

inspection is accomplished. The accident airplaneís most recent AS inspection

(an AS-1 inspection) was accomplished on July 13, 1996, at JFK, at an airplane

total flight time of 93,253 hours. 

ï TCS inspectionóincludes, to a greater degree and depth than the AS

inspection, the following: the cockpit and cabin; exterior fuselage, wings, gear,

and empennage; engines and struts; oil, hydraulic fluid, and water servicing;

and fuel tank sumping. TCS inspections are accomplished at intervals not to

exceed 1,200 hours of airplane time in service. The accident airplaneís most

recent TCS was accomplished on April 15, 1996, at JFK, at an airplane total

flight time of 92,231 hours. 

ï Station service (SS) inspectionóinterim cabin refurbishment program with

limited external airplane inspection. SS inspections are accomplished at

intervals not to exceed 800 hours of airplane time in service. The accident

airplaneís most recent SS inspection was accomplished on May 11, 1996, at

JFK, at an airplane total flight time of 92,466 hours.88

ï C inspectionóincludes checks of the integrity and airworthiness of the

airframe, fluid quantities, security of components, operational checks, overhaul

of specified components, the accomplishment of principal structure elements

per the structural inspection program, and the Corrosion Prevention and

Control Program. The C check consists of three blocks designated as 1E, 2E,

and 3E, accomplished in sequence at intervals not to exceed 13 months. The

accident airplaneís most recent C inspection (3E) was accomplished on

November 6, 1995, at an airplane total flight time of 91,084 hours. 

88 Although the SS inspection was to be accomplished at intervals not to exceed 800 hours of airplane

time in service, on July 12, 1996, the FAA had granted TWA a one-time 100-hour extension to this interval

for the accident airplane. Thus, the accident airplaneís next SS was due within 900 hours of the previous SS

inspection. 
Factual Information 45 Aircraft Accident Report

ï D inspectionóa more in-depth inspection of the structure and systems than the

C inspection to ensure the integrity and airworthiness of the airframe, fluid

quantities, security of components, operational checks, and overhaul of

specified components. D inspections are accomplished at intervals not to

exceed 4 years. The accident airplaneís most recent D inspection was

accomplished on December 14, 1992, at an airplane total flight time of

80,267 hours. The routine work is accomplished as defined in TWAís

Operations Specifications Manual, TWAís approved engineering reports,

structural inspection programs and documents approved for control of repair,

overhaul, parts replacement, periodic inspections, and routine checks.

1.6.3.1.1 TWAís General Inspection Policies

According to the TWA General Policies and Procedures Manual, dated May 15,

1994, TWA had the following three levels of airplane and powerplant inspection:

A. Visual Inspection: Visual airframe and power plant inspection constitutes a check of

visible or exposed areas, usually external, as specified on the appropriate inspection

forms. A visual inspection may include those items which are partially hidden, plus

those that might be readily accessible through quick access panels. 

B. Detailed Inspection: A detailed inspection is covered by two complementary

inspection conceptsóthe area concept and the specific item concept. 

Area ConceptóThe area inspection concept constitutes a very detailed

inspection of the designated area, including, but not limited to structures,

tubing, cables, wiring and any units exposed or visible through routine

open up. Normal assistance to visual inspection will be used as required

and may consist of mirrors, magnifying glasses, dye penetrant checks or

specialized non-destructive test equipment where applicable. 

Specific Item ConceptóThe specific item concept is a very detailed

inspection of a specific item as detailed on the work forms or by

inspection supervision. It is limited to the defined item(s) and does not

cover the associated area. 

C. Final Inspection: Upon completion of all maintenance and service work at

[C inspection] and higher maintenance as outlined on the appropriate maintenance

work forms, the airframe and powerplant shall be given a final check by an inspector. 

The airframe final inspection is intended as a visual safety check to ensure that all

access covers, inspection doors, and panels are installed; that tools, rags, loose hardware,

etc., have been removed from the cabin, cockpit, cargo compartments, wheel wells, and

engine inlets; and that all loose equipment and furnishings have been properly stowed.

The Safety Boardís review of TWA maintenance documents indicated that

inspection of the 747 CWT was to be accomplished by an area concept-type of detailed

inspection. The manual did not specifically define the area to be inspected, nor did it

provide specific inspection instructions for the various compartments of the CWT or the

various components within the CWT, such as FQIS wiring, connectors, and probes.
Factual Information 46 Aircraft Accident Report

1.6.3.1.2 TWA and Boeing Wiring Inspection Guidance

The Safety Boardís Maintenance Records Group asked TWA and Boeing

personnel to describe any procedures and guidance that they provided to maintenance

personnel regarding the protection of wire bundles in 747s while maintenance was being

accomplished near them. Boeing had published general inspection guidelines, which

advised operators to assess the existing condition of wires and use good judgment and

common sense during maintenance to prevent injury and damage. Boeingís guidance

indicated that wire bundles and equipment should be moved to avoid unnecessary damage

when any type of repair is being accomplished in the area. According to Boeing Service

Letter (SL) 747-SL-20-048, dated January 25, 1995, ìAs a general rule, wiring that is left

undisturbed will have less degradation than wiring that is reworked. As wiring and

components become more brittle with age this effect becomes more pronounced.î Boeing

had also issued Service Bulletin (SB) 747-53-2272, ìFuselageóNose Section 41 Body

Frame Structural Replacement and Reinforcement,î which advised operators to

ì[f]abricate a suitable work platform as required to protect wire bundles.î According to

Boeing engineers,89 most operators cover nearby wire bundles with bubble wrap or tape a

clean cloth over the bundle while making repairs in the area.

At the Safety Boardís request, TWA provided the Maintenance Records Group

with a copy of its Boeing Commercial Jet Standard Overhaul Practices Manual (most

recently revised in September 1997), which TWA indicated it had developed based on

Boeing documents. Chapter 20 of this manual included guidance regarding the following:

repair and replacement procedures for metal and electrical work, inspection procedures,

cleaning procedures, finishing procedures, installation procedures, specifications, and

materials. According to TWA, these subjects were also addressed in its maintenance

classroom training sessions, and maintenance/fleet service personnel were expected to use

general housekeeping procedures in performing their duties. TWA personnel indicated

that the company did not have a written policy concerning the use of protective covers

over or around wire bundles when performing metal work near the wire bundles; instead,

maintenance personnel were expected to determine (on a case-by-case basis) whether a

protective covering was necessary and whether the wire bundle needed to be moved to

accomplish a maintenance task. 

1.6.3.2 Accident Airplaneís Maintenance Information

Investigators reviewed the following records for the accident airplane: aircraft

maintenance logs for the accident airplane from December 1992 to the accident date;

C inspection routine and nonroutine work cards from June 3, 1986, to November 6, 1995

(the most recent C inspection); D inspection routine and nonroutine work cards from

February 8, 1986, to December 11, 1992 (the most recent D inspection);90 engineering

maintenance liaison records; and all records regarding airworthiness directives (AD), SBs,

and modification orders (MO)91 for the accident airplane. TWA maintenance records

indicated that the company had complied with all applicable FAA ADs,92 accomplished all

89 Similar information was also obtained in writing from TWA personnel. 
Factual Information 47 Aircraft Accident Report

scheduled maintenance items, and maintained its continuous airworthiness maintenance

program on the accident airplane. 

The Safety Boardís examination of the accident airplaneís maintenance records

revealed that the airplane had experienced several intermittent problems during fueling

operations in the 2 years before the accident, including one that occurred before the

accident airplane departed JFK. After the accident, a TWA mechanic advised the Board

that while the accident airplane was being fueled at JFK for TWA flight 800, the fuel

system shut down. The mechanic reported that the VSO fuse and an overflow circuit

breaker93 were pulled and the pressure fueling process was continued. After the airplane

was fueled, the mechanic reset the fuse and circuit breaker.94

In addition to the refueling difficulties noted, the Safety Boardís review of the

accident airplaneís maintenance records revealed three logbook entries regarding fuel

leaks during the preceding 2 years. The Board also noted 25 maintenance logbook entries

regarding fuel flow; fuel gauge indications, inaccuracies, and fluctuations; and inoperable

fuel system equipment. Appendix E contains descriptions of the fueling discrepancies

observed and the resultant maintenance actions. 

1.6.3.2.1 Airworthiness Directives and Service Bulletins Applicable to 747 

Fuel Pumps and Related Wiring

FAA AD 79-06-02, effective April 19, 1979, required compliance with Boeing

Alert Service Bulletin (ASB) 747-28-A2092, dated February 12, 1979, which

recommended a one-time main fuel tank pump wiring inspection, repair, and modification

to preclude electrical arcing95 into the Nos. 2 and 3 main fuel tanks.96 The AD described a

90 According to TWA personnel, in accordance with the companyís FAA-accepted maintenance

recordkeeping program, TWA retains its aircraft maintenance logs for 30 months, its C inspection routine

and nonroutine work cards for 10 years, and its D inspection routine and nonroutine work cards for 25 years.

(The FAA requires operators to retain inspection records until the next similar inspection or higher-level

inspection occurs.) 

91 MOs authorize the expenditure of funds and accomplishment of physical changes to an airplane or

components, one-time inspections and their resulting rework, and special scheduled work beyond normal

overhaul or maintenance. They also authorize funds for associated tooling, outside services, and initial

spares. MOs may be used to authorize service evaluations for the improvements in performance of TWA

airplanes, installation of new equipment, and extension of service life of airplanes and engine accessories

and components. 

92 See section 1.6.3.2.1 for information regarding ADs and SBs for the 747 fuel pumps and FQIS wiring

and section 1.6.3.2.2 for information regarding ADs and SBs for 747 structural inspections. 

93 A circuit breaker is a mechanical device designed to open a circuit (stop the current flow) when the

current flow within that circuit exceeds a set limit for a length of time. 

94 The TWA 747 Maintenance Manual states that if dispatch timing precludes repair of a malfunctioning

VSO control unit, ìthe shutoff system can be disabled by removing the fuse in the VSO control unit in the

right main equipment center on shelf E3. This permits operation of the refuel valves and fuel quantity

indicators, but disables the automatic VSO control unit. The fuel level shutoff must be controlled by the

individual refuel valve switches for all tanks.î 

95 Arcing is defined by Underwriters Laboratories, Inc. (UL), as ìa luminous discharge of electricity

across an insulating medium.î The electrical discharge of an arc can involve temperatures of several

thousand degrees Celsius. For additional information regarding arcing, see section 1.16.6.1. 
Factual Information 48 Aircraft Accident Report

method of alleviating wire chafing that involved installing a double layer of Teflon

sleeves97 over the wiring where it was routed within the fuel tanks. TWA maintenance

records indicated that the AD was accomplished on the accident airplane on June 6,

1979.98

On August 3, 1995, Boeing issued ASB 747-28-A2194, ìFuel-Distribution-Fuel

Boost and Override/Jettison Pumps-Inspection.î The ASB stated that operators had

removed eight fuel pumps from service that had fuel leaks at the pump/wire bundle

interface and recommended that all operators test the resistance on each 747 boost and

jettison/override pump ìat the next opportunity,î replacing those that did not pass the

insulation check. The SB stated the following: 

after a long time, water can get inside the potting of the wire terminal assembly

and cause corrosion. The corrosion in the wire terminal assembly can cause arcing

between the power pins and the pump case. The arcing causes thermal expansion

of the materials inside the cap. This expansion causes failure of the cap

attachment flange or the attaching screws and a subsequent fuel leak. 

According to TWA maintenance personnel, at the time of the accident,

ASB 747-28-A2194 had not been accomplished on the accident airplane because the FAA

had subsequently issued Notice of Proposed Rulemaking (NPRM) 96-NM-57-AD, which

proposed an AD that would ìrequire repetitive visual inspections to detect discrepancies

of the wire terminal assembly, electrical connector, and wire insulation on the fuel pump;

and replacement of the fuel pump with a new fuel pump, if necessaryî and ìrepetitive

insulation resistance test of the fuel pump wiring.î The resultant AD (AD 97-03-17)

became effective on March 14, 1997, and TWAís corresponding MO (72F57) was

finalized on April 1, 1997. 

1.6.3.2.2 Airworthiness Directives and Service Bulletins Applicable to 747 

Structural Inspections

Nose Section (Section 41)óInternal Inspections

On January 31, 1986, the FAA issued AD 86-03-51, which required structural

inspection of the nose section (section 41) on certain 747 airplanes, including the accident

airplane, ìto prevent sudden decompression of the fuselage.î On February 14, 1986,

Boeing issued SB 747-53A2265, ìFuselage-Nose Section 41-Body Frame Structure

Inspections and Crack Repairs,î which provided additional information regarding

inspection and repair of this area.99 According to the SB, ìnumerous body frame and other

internal structure cracks in the upper deck, main deck and lower lobe of the [nose section],

96 These fuel pumps are located in the outboard end of fuel tank Nos. 2 and 3 and draw fuel from the

lowest points of fuel tank Nos. 1 and 4. 

97 A sleeve is a woven or flexible jacket that protects electrical wiring. 

98 AD 79-06-02 (and postaccident superseding ADs) are discussed further in section 1.18.3.4. 

99 At the time of the accident, eight revisions to SB 747-53A2265 had been issued. The most recent

revision was issued on April 14, 1994, and specified that all frame cracks must be repaired before further

flight. 
Factual Information 49 Aircraft Accident Report

were reported by several operators on a number of airplanes with 7,000 to 20,002 flight

cycles....The body internal structure cracking is attributed to cabin pressure cyclic

loading.î The SB further stated that continued operation with undetected or unrepaired

frame cracks ìmay cause two or more adjacent frames to crack through near the same

stringer, which could lead to extensive body skin cracking and rapid cabin

decompression.î Therefore, the SB recommended inspections of the affected area to be

accomplished within specified flight cycle limits, with repairs as needed, and repeat

inspections at intervals not to exceed 3,000 flight cycles.100 The Safety Boardís review of

TWAís maintenance records indicated that the accident airplane was inspected in

accordance with SB 747-53A2265 on May 8, 1986. 

On December 15, 1986, the FAA issued AD 86-23-06, which superseded

AD 86-03-51 and required repetitive inspections of several areas in the 747 nose section.

TWA maintenance records indicated that the company accomplished the resulting initial

inspection on the accident airplane on April 15, 1988. On June 24, 1991, the FAA issued

AD 91-11-01, which superseded AD 86-23-06 and required several repetitive inspections

in accordance with Boeing SB 747-53A2265 to detect and repair body frame and other

internal structure cracks and possible adjacent skin cracks in the 747 nose section.

AD 91-11-01 further stated, ìinstallation of new and improved body frame structure in

accordance with FAA approved procedures or Boeing SB 747-53-2272, dated January 12,

1987Öis considered terminating action for the repetitive inspections required by this AD

for the structure (considered to be stringers, clips and skin associated with the frame).î

Boeing SB 747-53-2272, ìFuselage-Nose Section 41-Body Frame Structure Replacement

and Reinforcement,î stated, in part, ìfabricate a suitable work platform as required to

protect any wire bundles disconnected in the upper deck area between STA 420 and 520.î 

According to TWA maintenance records, an MO was written to direct and

schedule accomplishment of SB 747-53-2272. In accordance with this MO, the work

platform structure was fabricated on 17 of TWAís 747s, including the accident airplane,

terminating the need for the repetitive inspections required by AD 91-11-01. The MO also

directed modifications in accordance with Boeingís Document D6-35999, ìAging

Airplane SB Structural Modification Program-Model 747,î dated March 1989, which

required replacement of trailing edge flap tracks, reinforcement of the APU cutout,

installation of a splice strap at the STA 1241 bulkhead, and modification of longitudinal

floor beams. 

Sections 42 and 46 Lower LobeóInternal Inspections

On September 17, 1986, the FAA issued AD 86-18-01, requiring 747 operators to

inspect STAs 540 to 760 and 1820 to 1900 and stringers S-35 to S-42 on the left side of the

airplane ìto detect cracking of body frame structure in the lower lobe (sections 42 and 46)

of the fuselage.î AD 86-18-01 applied, in part, to all 747 series airplanes listed in Boeing

ASB 747-53A2237, Revision 1, dated March 28, 1986, which included the accident

airplane. According to TWA maintenance records, the accident airplane was inspected in

100 The SB indicated that operators should reduce the repeat inspection interval to 2,000 flight cycles on

airplanes with more than 20,000 flight cycles. 
Factual Information 50 Aircraft Accident Report

accordance with AD 86-18-01 on November 19, 1991, and the next repetitive inspection

was due on November 19, 1996. 

Nose Section (Section 41)óExternal Inspections

On January 22, 1991, the FAA issued AD 90-26-10, requiring 747 operators to

conduct external inspections of the fuselage skin from STA 220 to 520 and between

stringers S-6 and S-14 on both sides of the airplane ìto prevent rapid decompression of the

airplane.î Inspections were to be accomplished in accordance with Boeing

ASB 747-53A2321, dated October 31, 1989, and repeated thereafter at intervals not to

exceed 2,000 flight cycles. The AD required detected cracks to be repaired ìprior to

further flight.î According to TWA maintenance records, the accident airplane was

inspected in accordance with AD 90-26-10 on November 6, 1995, and the next repetitive

inspection was due on October 20, 1998. 

1.6.3.2.3 Maintenance Accomplished Near Fuel Quantity Indication System 

Wiring in the Accident Airplane 

The Safety Boardís review of TWAís maintenance records for the accident airplane

revealed several instances in which repairs and other maintenance items were

accomplished near the FQIS wiring paths during the 10-year period that preceded the

accident, including the following: 

ï In February 1986, during an inspection, a new forward cargo compartment

fitting was installed between the upper front spar chord and the cabin floor at

STA 1000, LBL 9. In addition, maintenance personnel cleaned the ìexposed

areas of all E1, E2, and E3 equipment shelves, wire bundles and terminal

stripsî and the wiring outboard of (behind) the flight engineerís panel in the

cockpit and removed metal shavings from the interior of the P-6 panel and the

right-side body landing gear wheel well (STAs 1350 to 1394). (The fuel flow

computer and VSO unit were located in the E3 equipment rack.)

ï In May 1986, during escape light path modification maintenance, a clamp was

installed on the right sidewall just above the main deck floor level in the

forward E/E compartment, securing a wire bundle, and a crack in the left-side

bulkhead at STA 440 was repaired. (The wire bundle that contained the fuel

flow computer and VSO unit wires was routed down the side of the airplane

from the flight engineerís position along STA 360, then turned aft beneath the

floor, passing through this area.) 

ï In April 1987, during an inspection, TWA maintenance personnel replaced two

cracked stringers and repaired a cracked former flange between STAs 420 and

460. Additionally, maintenance records from this visit indicated that debris

(including dirt, metal shavings/cuttings, and loose hardware) was cleaned from

the E/E compartments, and several loose/unclamped wires/wire bundles were

secured. (As previously indicated, the fuel flow computer and VSO unit were

located in the E3 equipment rack, and associated FQIS-related wiring is routed

through this area, between E3 and STA 360.) 
Factual Information 51 Aircraft Accident Report

ï In September 1988, during a C inspection, TWA maintenance personnel

observed evidence of electrical arcing at the cannon plug for the service carts

in galley C; two electrical cords were removed and replaced. Galley C is

located in the center of the airplane, just aft of the L2/R2 main cabin doors,

between STAs 850 and 880. Figure 20 is a diagram of the 747-100 interior.

(The CWT FQIS wire bundles [W480] were routed in a raceway beneath the

floorboards slightly aft of galley C.) 

ï In November 1988, during a D inspection, TWA maintenance personnel

removed numerous floor and wall panels located between STAs 360 and 1265

so that ìmetal workî repairs could be accomplished. (The areas addressed

during these repairs contained FQIS-related wire bundles that are routed

forward from the CWT rear spar, past the aft edge of galley C, to the fuel flow

computer and VSO unit.) 

ï In October 1990, during an inspection, TWA maintenance personnel observed

corrosion at STA 1241, LBL 57, removed the corroded metal, and repaired the

area. Subsequently, TWA maintenance personnel reinstalled wire bundles and

the wire bundle tray (which had been removed to gain access to the canted

pressure bulkhead) under the floor at STA 1241, LBL 70, after ìmetal workî

was completed. (W480 FQIS wires passed above and beneath this repair area.) 

ï In December 1991, during an inspection, TWA maintenance personnel

repositioned the wiring on the rear of the cockpit fuel quantity panel. (These

wire bundles included both power wires and FQIS circuit wires from each fuel

tank, including the CWT.) 

ï In December 1992, during a D inspection, the last of several tasks required by

a 747 structural modification program mandated by AD 90-06-06 was

accomplished. These tasks generated several nonroutine work cards, which

indicated that corrosion and cracks were found, and repairs were accomplished

in accordance with maintenance manual instructions. Several of these work

cards indicated that wire bundles and ducting were removed to gain access to

repair areas and reinstalled upon completion of repairs. Numerous wire bundle

clamps were found deteriorated and were replaced. At STA 920, a wire bundle

that was chafing was repositioned. (The work cards were not specific as to

which wire bundles were removed. The CWT FQIS wire bundle [W480] is

routed laterally along the floorbeam101 at STA 920 [crossing wires in the

raceway], en route to the cockpit.) 

ï On May 24, 1994, during scheduled ground maintenance at JFK (to perform a

cabin seating reconfiguration), the left potable water bottle cap burst and

separated from the cap attachment ring, resulting in damage to the area above

and around the potable water bottles. The CWT FQIS wire bundle was routed

through the area just outboard of the potable water compressor. According to

maintenance records, corrective actions included the following:

101 Although the floorbeams above the WCS are oriented longitudinally, floorboards in the rest of the

airplane are oriented laterally. 
Factual Information 52 Aircraft Accident Report

The potable water bottle air compressor pressure switch and connector,

water system pressure relief valve, plus 3 feet of associated wiring were

removed from the airplane for testing.102 The air compressor pressure

switch was reinstalled, and the connector and relief valve were replaced.

Both potable water bottles were removed, and the area, including the

bulkhead aft of the bottles, was inspected. Most of the potable water

bottlesí support rods and fittings were broken and/or bent and were

replaced or repaired. The potable water bottles were replaced with

serviceable potable water bottles. 

Several damaged ceiling and floor panels, baggage restraint bars,

passenger service units, and trim panels were observed near seat rows 19

through 22, seats 4, 5, 6, and 7, and were removed and replaced. At

STA 980, the floor beam above the left potable water bottle was removed,

repaired, and reinstalled. The ceiling support in the forward cargo

compartment was bent at LBL 24 between STAs 960 and 980, and the

floor support was bent and cracked at STA 984, between LBL 12 and

LBL 32; both supports were removed, repaired, and reinstalled. 

Wire bundles forward and aft of STA 970 were removed from the support

clamps for sheet metal repair (forward and above the potable water bottle

area) and were resecured upon repair completion. The wire bundle for the

potable water bottle air compressor was disconnected just forward of the

compressor during repairs, and was reconnected and secured after the

metal repairs were completed. 

The forward cargo compartment between STAs 980 and 1000 was

cleaned of water, dirt, and debris.

ï In November 1995, TWA maintenance personnel accomplished AD 89-12-07,

ìOverwing Center Section Cavity Drain Inspection,î during which the main

cabin floor panels were removed to expose the STA 1240 cavity drain. (The

FQIS wire bundle that was routed from the cockpit to the CWT is located

beneath the floorboards near this work.) 

During its examination of recovered wreckage, the Safety Board found evidence of

repairs accomplished near FQIS wire routing areas for which no associated maintenance

records were found. For example, wires carrying FQIS signals to the upper deck AIDS

unit were routed across the top of the airplane through an area where replacement

fasteners and differences in paint were observed. The interior of the airplane skin in this

area had a note, handwritten in green paint, that stated, ìOCT 1, 1992 1:47 AM.î The

Maintenance Records Group did not find a maintenance record that it could associate with

this handwritten note. 

102 As a result of this potable water bottle event and subsequent tests, on May 31, 1994, Boeing advised

TWA to replace the tested relief valve with a different part number (P/N) pressure relief valve. On

October 27, 1994, Boeing issued SB 747-38A2105, which advised all 747 operators (line numbers 1 through

1,013) to replace the potable water system pressure relief valve at the earliest opportunity. On June 23, 1995,

the FAA issued AD 95-11-03, which required replacement of the potable water system pressure relief valve. 
Figure 20. A diagram of the 747-100 interior.
Factual Information 54 Aircraft Accident Report

1.6.3.2.4 TWA Flight 800 Predeparture Maintenance Information

The flight and cabin crew of TWA flight 881 (the accident airplaneís inbound

flight segment, from Athens, Greece, to JFK) entered several maintenance writeups in the

logbook, which were addressed by TWA maintenance personnel at JFK before TWA flight

800 departed. These writeups included the following:

ï The Nos. 1 and 2 engines exceeded the exhaust gas temperature limits of

925° F for 2 seconds. The overtemperature was caused by an unanticipated

rapid spool-up of the engines during takeoff out of Athens. Maintenance

personnel checked the inlet and exhaust on both engines, and no abnormalities

were found. 

ï A deferred item from July 7, 1996, indicated that maintenance personnel had to

physically check the No. 3 engine oil quantity because of an oil quantity gauge

discrepancy. The No. 3 engine oil tank was checked and serviced with oil. No

malfunctions were found. 

ï The splash guard in the aft upper deck lavatory was missing. Maintenance

personnel reset the splash pan. 

ï The drain in galley C had leaked, and the surrounding area of the floor was

soaked. Maintenance personnel cleared an obstruction from the drain.

ï The No. 1 engine pressure ratio (EPR)103 indicator was sticking. Maintenance

personnel replaced the EPR module. 

As previously indicated, maintenance personnel also accomplished a PS inspection

and checked the landing gear tire pressure before the accident airplane departed JFK. The

most recent TWA All Open Item Work Sheet for the accident airplane indicated that the

following maintenance items had been deferred: 

ï A June 9, 1996, maintenance logbook entry indicated that a 30-inch crack was

observed in the fiberglass panel on the underside of the left wing, between the

fuselage and the wing landing gear at the trailing edge, forward of the flap.

Maintenance personnel performed a repair;104 however, the item remained

open until a new panel could be produced and installed. 

ï A June 30, 1996, maintenance logbook entry indicated that a tear was observed

in the arm rest cover at row 31, seat 7. The item was deferred with a note that

stated that the entire seat should be recovered.

ï A July 4, 1996, maintenance logbook entry indicated that the rear attachment

on the left-side canoe fairing for the No. 2 trailing edge flap was broken. The

fairing was removed, flap carriage was reported to be otherwise normal, and

103 EPR is a measurement of the engineís power output as a ratio of total pressure of the gases in the

exhaust pipe divided by the total pressure of the air entering the engine inlet. 

104 This time-limited repair would typically have been considered an airworthy repair until the airplaneís

next scheduled higher maintenance check; in the case of the accident airplane, this would have been a D

check scheduled for December 1996/January 1997. However, TWA had scheduled the panel for replacement

when the new panel was available, at the next SS inspection. 
Factual Information 55 Aircraft Accident Report

replacement of the fairing was deferred on the basis of the Minimum

Equipment and Dispatch Procedures, Configuration Deviation List. The

following operational stipulations were associated with this deferment: (1) for

takeoff, reduce runway zero wind and climb limit weights 2,700 pounds;

(2) for landing, reduce critical temperature 1° F. A placard listing these

limitations was affixed in the cockpit. 

ï A July 5, 1996, maintenance logbook entry described spots on the upper deck

carpets. The item was deferred because there was insufficient time to perform

the corrective action.

ï A July 7, 1996, maintenance logbook entry indicated that the No. 3 engine

failed to go into reverse thrust and that the thrust reverser lever interlock failed

to release. Maintenance personnel were unable to complete the repair because

they did not have the necessary cables in stock. The item was deferred until

July 17, 1996, when a mechanic was scheduled to install the cables while the

airplane was on the ground at JFK. However, the mechanic told Safety Board

investigators that he was unable to complete the repair before the airplaneís

scheduled departure, so he mechanically locked the thrust reverser in the

forward position, and the item was again deferred for 10 days. The Minimum

Equipment and Dispatch Procedures indicated that the flight crew must be

notified, associated systems for stopping the airplane must be operational, and

affected systems must be placarded ìINOP.î Records indicated that these

procedures were followed. 

ï A July 7, 1996, maintenance logbook entry indicated that maintenance

personnel replenished the oil supply in the No. 3 engine, after which the flight

engineerís oil quantity gauge indicated that the engine contained 4 quarts of

oil. After the engine was started, the flight engineerís No. 3 oil quantity gauge

indicated 2.3 quarts. The flight engineerís oil quantity gauge was placarded

ìINOP,î and the item was deferred. TWAís minimum equipment and dispatch

procedures indicate that the items must be repaired within 10 calendar days.

The procedures state that ìone (oil quantity gauge) may be inoperative if it is

verified before each takeoff that the oil tank is filled to the maximum

recommended capacity; there is no evidence of above normal oil consumption

or leakage; and oil pressure indicating, low oil pressure warning, and oil

temperature indicating systems operate normally and are monitored.î The

affected gauge (flight engineerís panel) was to be placarded ìINOP.î Before

the departure of TWA flight 800, these procedures were signed off as

accomplished.

ï A July 11, 1996, maintenance logbook entry indicated that the drain in

galley C was clogged and drained very slowly. Maintenance personnel cleared

the drain and signed off the maintenance item. However, a July 12, 1996,

maintenance logbook entry indicated that the drain was still clogged.

Maintenance personnel again cleared the drain, blew air through the line, and

signed off the maintenance item. Later that day, at JFK, a logbook entry

indicated that the drain was still clogged; however, because maintenance

personnel did not have time to address the item, it was deferred. 
Factual Information 56 Aircraft Accident Report

ï A July 15, 1996, maintenance logbook entry indicated that the No. 3 left

leading edge flap amber (cautionary) light illuminated when the leading edge

flaps were in the retracted position. The item was deferred because

maintenance personnel did not have time to complete the task. The Minimum

Equipment and Dispatch Procedures indicated that the affected lights must be

placarded ìINOPî and that the item must be repaired within 10 calendar days.

ï A July 15, 1996, maintenance logbook entry indicated that the R3 door

emergency chute pressure gauge was not visible. Maintenance personnel

deferred the item, indicating that the emergency chute cover needed to be

ìreworked.î 

ï A July 17, 1996, maintenance logbook entry indicated that the captainís

weather radar display was inoperative, indicating no range markings or

antenna sweeps. The item was deferred. The Minimum Equipment and

Dispatch Procedures indicated that the indicator must be placarded ìINOPî

and that the item must be repaired within 10 calendar days. 

1.7 Meteorological Information 

The surface weather observation taken about 1951 at JFK on July 17, 1996,

located 52 nautical miles (nm) west of the accident site, stated the following: 

Winds 220° at 8 knots; visibility 10 miles; clouds at 6,000 feet scattered, ceiling

7,000 feet broken, 9,000 feet broken; temperature 82° F; dew point 70° F;

altimeter setting 30.07 inches of Hg [mercury]; wind shift 1906; rain began 1918

and ended 1929; precipitation 0.00 inch between 1927 and 1951. 

The surface weather observation taken about 2051 at JFK stated the following:

Winds 240° at 9 knots; visibility 10 miles; few clouds at 10,000 feet; temperature

80° F; dew point 69° F; altimeter setting 30.08 inches of Hg. 

The surface weather observation taken about 1945 at Francis S. Gabreski Airport

(FOK) Westhampton Beach, New York, located 12 nm north of the accident site (the

nearest reporting station to the accident site), stated the following:

Winds 240° at 4 knots; visibility 4 miles; haze; clouds at 6,000 feet scattered;

temperature 73° F; dew point 66° F; altimeter setting 30.08 inches of Hg; total sky

cover 3/8.

The surface weather observation taken about 2045 at FOK stated the following:

Winds calm; visibility 6 miles; haze; clouds at 6,000 feet scattered; temperature

72° F; dew point 66° F; altimeter setting 30.09 inches of Hg; total sky cover 3/8.

The winds aloft measured by a weather balloon launched from Upton, New York,

on July 17, 1996, about 2000, are shown in table 3. (Upton is located about 15 nm from

the accident site.) 
Factual Information 57 Aircraft Accident Report

These numbers were used during the Safety Boardís trajectory study.105 Review of

the meteorological data revealed no record of significant meteorological conditions in the

area or at the time of the accident. 

1.8 Aids to Navigation

No difficulties with the navigational aids were known or reported.

1.9 Communications

No difficulties with communications were known or reported.

Table 3. Winds aloft measured by a weather balloon launched 

from Upton, New York.

Altitude (feet msl)

Wind Direction 

(degrees) Wind Speed (knots)

1,000 270 12

2,000 280 14

3,000 285 17

4,000 290 17

5,000 303 19

6,000 310 19

7,000 315 17

8,000 320 16

9,000 330 12

10,000 335 12

11,000 320 12

12,000 295 16

13,000 290 16

14,000 300 17

15,000 303 19

16,000 305 21

17,000 315 29

18,000 315 33

105 For additional information regarding the trajectory study, see section 1.16.2.1.
Factual Information 58 Aircraft Accident Report

1.10 Airport Information

JFK is located 1/2 mile southeast of the New York City limits and has an airport

elevation of 13 feet. The airport has five runways. Runway 14/32 is 2,560 feet long and

75 feet wide; runway 13R/31L is 14,592 feet long and 150 feet wide; runway 13L/31R is

10,000 feet long and 150 feet wide; runway 4R/22L is 8,400 feet long and 150 feet wide;

and runway 4L/22R is 11,351 feet long and 150 feet wide. The accident airplane departed

from runway 22R. 

1.11 Flight Recorders

The two flight recorders on the accident airplane had been mounted in the aft

fuselage, above and aft of the L5 door. According to wiring diagrams and maintenance

documents and personnel, 115-volt a.c. power (provided by the engine-driven generators)

was routed to the recorders through wiring along the upper right side of the passenger

cabin from the a.c. essential bus/flight engineer panel. The recorder signal wires were

routed along the upper left side of the passenger cabin. The recorders were recovered from

the Atlantic Ocean by surface-supplied U.S. Navy divers operating from the U.S.S Grasp

on the evening of July 24, 1996. They were immediately packed in water to prevent/delay

the onset of corrosion and shipped to the Safety Boardís laboratory in Washington, D.C.,

for readout. 

1.11.1 Cockpit Voice Recorder

The CVR installed on the accident airplane was a Fairchild model A-100.106

Although the CVR unit exhibited external and internal structural damage and the

recording medium (magnetic tape) was wet (which, as noted in section 1.1, resulted in

sounds similar to recording tape damage noise to be recorded at 2031:05), the tape was

otherwise in good condition, and the quality of the recording was good.107 The CVR

recording consisted of four channels of audio information: one channel contained audio

information recorded by the CAM, and the other three channels contained audio

information recorded through the radio/intercom selector panels at the captain, first

officer, and flight engineer positions. The accident airplane was not equipped with

noise-activated (ìhotî) microphones at the flight crew positions (nor was it required to be

so equipped). Therefore, the audio information recorded at the flight crew positions

consisted of air-to-ground communication and navigation radio audio information and did

not include flight crew conversations and sounds in the cockpit. However, depending on

106 The CVR identification plate and S/N were missing and were not recovered.

107 The Safety Board uses the following categories to classify the levels of CVR recording quality:

excellent, good, fair, poor, and unusable. A good recording is one in which most of the crew conversations

can be accurately and easily understood. The transcript that is developed may indicate several words or

phrases that are not intelligible. Any loss in the transcript can be attributed to minor technical deficiencies or

momentary dropouts in the recording system or to a large number of simultaneous cockpit/radio

transmissions that obscure each other.
Factual Information 59 Aircraft Accident Report

their nature and volume, these sounds could be recorded by the CAM. The cessation of the

CVR recording at 2031:12 was consistent with the loss of electrical power to the recorder.

A transcript was prepared of the entire 31-minute 30-second recording. See appendix B

for a complete transcript of the CVR recording. 

1.11.1.1 Sound Spectrum Study 

The Safety Board further examined the audio information recorded by the four

CVR channels using a sound spectrum analyzer (which provides a visual presentation of

the frequency of the sound signals) and a computer signal analyzer (which allows analysis

of the analog wave form and frequency content of the sounds and provides detailed timing

information of the events).108 Examination of the CVR sound spectrum information

indicated that throughout most of the accident flight, all of the CVR flight crew position

channels recorded an electrical background noise that consisted primarily of a 400-Hz

frequency electrical system power ìhum,î with harmonic tones109 at multiples of up to

13 times the 400-Hz frequency.110 However, 0.73 and 0.68 seconds before the CVR

recording stopped, there were brief (2 microseconds) changes in the electrical system

background noise hum recorded by the captainís position CVR channel. The harmonic

tones of multiples greater than 800 Hz were not recorded (were ìdropped outî), whereas

the 400-Hz electrical background noise was still recorded on the captainís position

channel. There was no evidence of any other power disturbances on any channel of the

CVR recording.

The Safety Board contracted with the Naval Air Warfare Center-Aircraft Division

(NAWC-AD), Patuxent River, Maryland, to measure the voltage harmonics of the

captainís CVR channel under various electrical load conditions.111 According to the

NAWC-AD report,112 the baseline total harmonic distortion113 of the voltages measured at

the captainís CVR channel input was approximately 33 percent. The NAWC-AD testing

indicated that application of an electrical load resulted in a reduction in the total harmonic

distortion voltages measured at the captainís CVR channel. The electrical load adjustment

108 For additional details and graphs, see Sound Spectrum Group Chairmanís Factual Report, dated

November 15, 1997. 

109 Most noises, including electrically generated tones, are made up of a fundamental frequency (in this

case, 400 Hz) and multiples of that fundamental frequency, which are known as harmonic tones. 

110 The CVR wiring is routed from the cockpit to the tail of the airplane with numerous wires and cables

that are powered by the airplaneís 115-volt (400-Hz) electrical system, which results in the CVR recording a

400-Hz background noise. Most airplanesí a.c. electrical systems operate on a 400-Hz frequency; this

background noise hum is a common feature on CVRs. 

111 The Safety Board also contracted with the NAWC-AD to determine whether electrical power

transients could induce sufficient energy into the FQIS wiring through capacitive and inductive coupling to

ignite fuel in a 747-100ís CWT. For additional information regarding these tests, see section 1.16.6.4.2. 

112 See NAWC-AD Report No. NAWCADPAX/TR-2000/33, Boeing 747-100 Fuel Quantity Indication

System (FQIS) Susceptibility to Induced Energy from Capacitive and Inductive Cable Coupling, dated

June 12, 2000.

113 Total harmonic distortion is used to compare an electrical waveform to a pure sine-wave, expressed as

a percentage of pure sine-wave voltage output. A total harmonic distortion score of 0 percent is the

equivalent of a perfect sine-wave and indicates no harmonic distortion. 
Factual Information 60 Aircraft Accident Report

required to attain the minimum value of total harmonic distortion voltages observed in the

NAWC-AD tests (less than 12 percent) occurred under the following conditions:

immediately after closing all generator circuit breakers; when the circuit breaker for

TRU 2 was opened; when the essential power selector was switched to either electrical

bus 1, 2, or 3; and when all six TRU circuit breakers were opened. 

As previously discussed in section 1.1, the CVR recorded an event (a ìvery loud

soundî) that was about 40 percent louder than the previous signals during the last few

tenths of a second of the CVR recording, which continued until the CVR recording

abruptly stopped.114 This event was recorded by all four CVR channels; however,

examination revealed that the very loud sound was recorded by the captainís position

channel 7.25 microseconds before it was recorded by the other three channels. 

1.11.1.2 Cockpit Voice Recorder-Related Airplane Tests

The Safety Board documented the CVRís response to various types of explosive

events during a series of controlled tests on a decommissioned 747-100 airplane.115 The

test series included the following conditions: detonation of four simultaneous explosions

in four different cargo container locations on board the pressurized airplane, controlled

high explosive detonations at various locations inside and outside the unpressurized

fuselage, and 15 fuel/air mixture explosions at various locations on board the

unpressurized airplane. The fuel/air mixture explosion tests included one test in which a

known fuel/air mixture was exploded in the airplaneís CWT.116

Sound spectrum analysis plots from these airplane tests were compared with those

from the TWA flight 800 CVR recording. For further comparisons, the Safety Board

plotted the CVR recordings from other known in-flight explosions/breakups (such as

Pan Am flight 103, a 747-100 airplane that crashed at Lockerbie, Scotland, after a bomb

on board exploded;117 an Air India 747-100 that crashed in the Atlantic Ocean southwest

of Ireland after a bomb on board exploded;118 and United flight 811, a 747-100 that lost its

forward cargo door in flight.119 The Board also plotted the CVR recording from a

Philippine Airlines 737-300 that experienced a fuel/air mixture explosion in the CWT as it

114 This sound cannot be discerned simply by listening to the CVR but was identified through the sound

spectrum study; therefore, it is not indicated in the CVR transcript. 

115 This airplane was obtained by the FAA and the British Civil Aeronautic Administration to conduct

explosive hardening trials on cargo containers. The airplane fuselage was structurally intact, with all exterior

doors and windows in place; therefore, it could be pressurized. However, the airplaneís engines, the cabin

interior, and the cockpit instruments had been removed. For the Safety Boardís tests, the airplane was

equipped with additional instrumentation (including multiaxis accelerometers, pressure sensors, flash or

detonation sensors, cabin microphones, and CAMs) and several CVRs, one of which approximately

duplicated the CAM and recorders that were installed on TWA flight 800. 

116 For the CWT fuel/air mixture explosion test, additional instrumentation was added to record the

acceleration, pressures, and the gas mixture inside the tank. For additional information regarding these tests,

see section 1.16.5.6. 

117 See Air Accidents Investigation Branch. 1990. Report on the Accident to Boeing 747-121, N739PA at

Lockerbie, Dumfriesshire, Scotland, on 21 December 1988. Aircraft Accident Report 2/90.

118 See Report of the Court Investigating. February 26, 1986. Accident to Air India Boeing 747 Aircraft

VT-EFO, ìKanishkaî on 23 June 1985. Honorable Mr. Justice B. N. Kirpal, Judge, High Court of Delhi.
Factual Information 61 Aircraft Accident Report

was being pushed back from the gate at Ninoy Aquino International Airport, Manila,

Philippines, on May 11, 1990.120

The CVRs recovered from these airplanes all recorded very loud sound events just

before they stopped recording. The sound signatures from these events were compared

with the sound signatures recorded at the end of the TWA flight 800 CVR recording.

Generally, the sound signatures could be characterized based on how quickly the loud

noise event rose from the background noise (rise time), the duration of the loud noise

event, and how quickly the loud noise event decreased (fall time).121 The TWA flight 800

CVR recorded noise characteristics that were most similar to those recorded by the CVRs

on board the United flight 811 and Philippine Airlines airplanes. The loud noise events

recorded by these three CVRs were characterized by longer rise times, durations, and fall

times than the loud noises recorded by the CVRs on board the Pan Am and Air India

airplanes; the Pan Am and Air India CVR recordings exhibited very fast rise times, very

short durations, and very fast fall times. The TWA flight 800 CVR was the only CVR that

recorded the change in the airplaneís electrical system background noise described in

section 1.11.1.1.

1.11.2 Flight Data Recorder

The FDR was a Sundstrand model UFDR, S/N 10291, which recorded

18 parameters of airplane flight information on a 1/4-inch magnetic 8-track tape capable

of retaining 25 hours of data. The accident airplaneís FDR system included a data

acquisition unit, which gathered and converted analog flight information to digital data for

transmission to the central electronics unit (CEU). The CEU provided a serial binary data

stream to the FDR. The 18 recorded parameters included time (Greenwich mean time),

pressure altitude, indicated airspeed, magnetic heading, vertical acceleration, longitudinal

acceleration, angle-of-attack (AOA), pitch attitude, roll attitude, pitch trim (stabilizer)

position, aileron position (sensed at the right inboard aileron), elevator position (sensed on

right and left sides), rudder position (sensed for upper and lower panels), flap position

(sensed for inboard and outboard flaps on both wings), leading edge devices (sensed for

each deviceófour on each wing), EPR for all four engines, thrust reverser position for all

four engines, and VHF microphone keying. 

Although the FDR exhibited external and internal impact and water damage, the

crash enclosure and recording medium were intact and yielded data of good quality.

Examination of the recorded data indicated that although the FDR operated normally until

119 See National Transportation Safety Board. 1990. United Airlines Flight 811, Boeing 747-122,

N4713U, Honolulu, Hawaii, February 24, 1989. Aircraft Accident Report NTSB/AAR-90/01. Washington,

DC. 

120 The Safety Board assisted in the Philippine Governmentís investigation of this accident. Damage to

float switch wiring and a defective fuel quantity sensor were identified as possible sources of ignition;

however, a definitive ignition source was not confirmed. For information about the safety recommendations

that resulted from this investigation, see section 1.18.1. 

121 The amplitude of the noise was not a reliable means of comparison because these noise events

typically overloaded the CVR recording system when they occurred. 
Factual Information 62 Aircraft Accident Report

it stopped recording at the time of the power loss, some parameters recorded anomalous or

erratic values. The following parameters contained frequent data anomalies recorded

during the accident and previous flights: 

ï Aileron position ñ Recorded values were noisy and erratic.

ï Thrust reverser engine No. 3 ñ During the landing at JFK that preceded the

accident, the ìtransitî indication was displayed while the other three engines

indicated a ìdeployedî condition. (The thrust reverser on engine No. 3 had

been mechanically locked in the retracted position by maintenance personnel

at JFK before TWA flight 800 departed on the day of the accident. Corrective

maintenance was deferred for up to 10 days.) 

ï Altitude (coarse) ñ Indicated erroneous altitude values and was unusable. 

ï Altitude (fine) ñ Occasionally displayed noisy values.

ï Leading edge flap left No. 3 ñ With the flaps retracted, the ìtransitî indication

remained on. With the flaps extended, the ìtransitî condition indicated

ìextended.î The maintenance log contained the following entry: ìJuly 15th,

3L LE flap amber [light] stays on with LE flaps up and retracted electrically.

[Forward] panel lights ops check ok.î 

ï Indicated airspeed ñ The recorded values were occasionally erratic. 

Examination of the FDR data indicated that the data recorded during the accident

airplaneís approach and landing at JFK before the accident flight (as TWA flight 881) did

not reveal any anomalous airplane or flight conditions. During the first 12 1/2 minutes of

the accident flight (from the start of the takeoff roll until 2031:12, when the recording

stopped abruptly), the FDR operated continuously and recorded data consistent with a

normal departure and climb. The data indicated that the airplane was in a wings-level

climb, and the vertical and longitudinal acceleration forces were consistent with normal

airplane loads when the recording stopped. Examination of the FDR data revealed that the

interruption of the recording at 2031:12 was consistent with the loss of electrical power to

the recorder.122

1.12 Wreckage Recovery and Documentation 

Information

Pieces of the wreckage were distributed along a northeasterly123 path about 4 miles

long by 3 1/2 miles wide in the Atlantic Ocean off the coast of Long Island. (Figure 21

shows the wreckage location relative to the airplaneís flightpath, JFK, and Long Island.)

122 For additional information, see Flight Data Recorder Group Chairmanís Factual Report (Revision 1),

Flight Data Recorder Group Chairmanís Factual ReportóAddendum I, and Flight Data Recorder Group

Chairmanís Factual ReportóAddendum II, all of which are dated February 15, 2000. 

123 At the last point of radar contact, the airplaneís true course was about 75°; the wreckage was generally

oriented along a path of about 60° from that last point of radar contact. 
Factual Information 63 Aircraft Accident Report

The first priority of the early search and recovery efforts was the recovery of victims;

wreckage recovery was the second priority.124

Throughout most of the search and recovery operation, remote-operated vehicles

(ROV), SSS, and laser line-scanning equipment were used to search existing underwater

debris fields. This equipment was also used to explore areas outside existing debris fields

that might contain victims and wreckage and to videotape the wreckage and ocean floor.

Scuba divers and ROVs were used to recover victims and wreckage. In the late stages of

wreckage recovery, scallop trawlers were used to recover pieces of wreckage that had

become embedded in the ocean floor. (Search and recovery efforts are described in detail

in appendix C.) 

When pieces of wreckage were recovered, they were typically tagged and

numbered according to their wreckage recovery location. Although most tagged pieces of

wreckage were tagged on the ship as they were recovered, it occasionally became

necessary to cut or separate objects (previously tagged as a whole) into more than one

piece. Additionally, some objects were extracted from an entangled group of debris

(recovered and tagged as a unit); in some cases, pieces were received in a bag, net, or box

full of other items with one tag assigned to the container. Finally, some parts simply broke

during handling/transport, leaving some parts untagged. In all of these situations, the

recovery position information on the ship tag from the original object or group of objects

was transferred to the hangar tag(s) assigned to the separated object(s). The

documentation/tagging process is described in detail in appendix D.125

Pieces of wreckage were then typically transported by boat to shore through the

Shinnecock Inlet, where they were loaded onto trucks and transported to leased hangar

space at the former Grumman Aircraft facility in Calverton, New York.126 Upon arrival at

the hangar, investigators worked to identify and document the pieces of wreckage.

124 Very few pieces of wreckage were recovered during the first 2 days after the accident; those that were

recovered during this time were found floating on the oceanís surface. Several days of side-scan sonar (SSS)

searching preceded the underwater wreckage recovery operations. 

125 During the wreckage recovery phase, an extensive database of recovery information was created. This

database contained a variety of information collected during the search and recovery operation, including

recovery position data for thousands of parts. Extensive efforts involving representatives from the parties to

the investigation were undertaken to cross-check and validate the information in this database. Nonetheless,

not all of the recovery positions listed in the database are known with the same degree of certainty. A

detailed Data Management Report was prepared to fully document this issue and to ensure proper use of the

database. When used in accordance with the provisions of this report, the tags database provides a highly

accurate source of recovery data. Recovery position data from the tags database were used in the trajectory

study and were mentioned in the Sequencing Groupís study report. Safety Board investigators audited the

recovery data used in the trajectory and sequencing studies to confirm the quality of the data extracted from

the tags database. This validation work confirmed that the data were used appropriately in both of these

studies. For additional information regarding the tags database, see appendix D or the Data Management

Report, dated November 17, 1997. 

126 Although almost all of the recovered wreckage was transported by boat to shore through the

Shinnecock Inlet, during the first few days of recovery, some recovered wreckage was transported to shore

through the Moriches Inlet. Additionally, some pieces of wreckage were flown from the ships to the hangar

at Calverton. 
Figure 21. The wreckage location relative to the airplaineís flightpath, JFK, and Long Island.
Factual Information 65 Aircraft Accident Report

The recovery effort took more than 10 months and involved personnel and

equipment from multiple agencies and companies. Most of the examination,

documentation, and (where pertinent) reconstruction of the recovered pieces of wreckage

were completed within 1 year of the accident. However, some wreckage examination was

ongoing until mid-2000. 

More than 95 percent of the accident airplane wreckage was eventually recovered.

Throughout the wreckage recovery and documentation processes, fire and explosives

experts and/or metallurgists from the Safety Board, DoD, Federal Bureau of Investigation

(FBI), Bureau of Alcohol, Tobacco, and Firearms (ATF), FAA, and parties to the

investigation thoroughly examined all recovered pieces of the wreckage for evidence of

damage characteristic of a bomb,127 missile, or high-order explosive (such as hot-particle

penetration, pitting, petalling, hot-gas erosion of metal, high degree of fragmentation, hole

penetration from outside to inside).128 No evidence of such damage was found. Where

areas of small missing pieces of the airplane were observed in the reconstructions, pieces

of wreckage that would have been adjacent to or nearby those areas were reexamined

closely. These adjacent or nearby pieces of wreckage contained no evidence of damage

from a bomb, high-order explosive (bomb or missile warhead), or from missile warhead

entry or detonation. 

As previously noted, a few pieces of the airplane structure were recovered floating

in the water. The most prominent of these pieces was a very large portion of the right

wing, from just outboard of the outboard engine to the wing tip. However, most of the

pieces of airplane wreckage were recovered from the ocean floor, predominantly in three

identified debris fields labeled by investigators the red, yellow, and green zones, from

farthest west to farthest east, respectively. (Figure 22a shows the location of the three

debris zones and the Long Island shoreline. Figures 22b and 22c show a 747 and the

accident airplaneís flightpath, respectively, color-coded to indicate the debris fields from

which corresponding wreckage was recovered.) The following subsections provide a

general description of the pieces of wreckage (and their major characteristics) recovered

from each zone. Additionally, these subsections describe the recovered condition of the

accident airplaneís engines, CWT fuel pumps and other CWT components, the air

conditioning equipment, and the electrical components and wiring. Details of damage and

specific characteristics of the various structural pieces are given, as necessary, to support

the discussion of the breakup sequence in section 1.16.3. See appendix F for fire

damage129 and soot deposit diagrams for recovered pieces of wreckage from the wings, aft

fuselage, CWT, and other selected areas. For additional information, see Fire and

Explosion Group Chairmanís Factual Report, dated August 22, 1997. 

127 For the purposes of this report, the term ìbombî means an explosive device designed to release

destructive material at high velocity upon detonation, but does not include an explosive device designed

only to set off a small charge of sufficient strength to penetrate the fuel tank and ignite explosive vapors. 

128 Several factors led to speculation that the accident might have been caused by a bomb or missile

strike, including heightened safety and security concerns because of the 1996 Olympics then being held in

the United States, the fact that TWA flight 800 was an international flight, and the sudden and catastrophic

nature of the in-flight breakup. 

129 During the hours after the accident, there were numerous fuel-fed fires on the surface of the water

around debris in the green zone. Thus, some of the fire damage and soot deposits found on pieces of

wreckage discovered in the green zone may have occurred postimpact. 
Figure 22a. Map showing the locations of the red, yellow, and green zones. 
Figure 22b. A 747, color-coded to indicate the debris fields from which corresponding wreckage was recovered. 
Figure 22c. The accident airplaneís flightpath, color-coded to indicate the debris fields from which corresponwreckage was recovered. 
Factual Information 69 Aircraft Accident Report

1.12.1 Wreckage Recovered from the Red Zone

The red zone was the largest of the three zones and was located farthest west

(closest to JFK) in the wreckage distribution. Pieces recovered from the red zone generally

included pieces from between about fuselage STA 840 and about fuselage STA 1000 (the

aft portion of section 42ósee figure 3a for station references); the structure from the aft

end of the forward cargo compartment; and pieces from the WCS, including most of the

front spar, a large portion of SWB3, and the manufacturing access door from SWB2.

Other pieces of wreckage recovered from the red zone included three of the four nose

landing gear doors (the fourth nose landing gear door was located in the yellow zone),130

the forward portion of the keel beam, main cabin floor beams, flooring material from

above the WCS, galley C, several passenger seats, two cargo containers, and the two

forward air conditioning packs131 (packs 1 and 3ópack 2 was recovered from the green

zone).132

None of the pieces of wreckage recovered from the red zone exhibited crushing

damage133 as severe as that found on many pieces of the structure recovered from the

yellow and green zones.134 In addition, although some of the pieces recovered from the red

zone contained light soot deposits, none of the pieces contained moderate or heavy soot

deposits or any other evidence of exposure to fire. Specifically, no fire damage was

observed on the forward portion of the keel beam, air conditioning pack debris, and the

fairing material located beneath the CWT. 

Two of the four cargo containers that had been loaded in the accident airplaneís

forward cargo compartment (those located on the left side of the forward cargo

compartment) were recovered from the red zone. The sides of these two containers were

largely intact, and the damage observed was consistent with water impact. Examination of

the forward cargo compartment containers and recovered portions of the forward cargo

compartment and forward cargo compartment equipment (including the potable water

bottles, floor tracks, cargo container stops)135 revealed no evidence of preimpact damage

to the forward cargo compartment ceiling and/or wiring routed along the ceiling.136 There

was no evidence that the potable water bottles had contacted the cargo containers, and all

130 For more information about the yellow zone, see section 1.12.2.

131 For specific descriptions of the air conditioning packs, see section 1.12.6.

132 According to wreckage recovery documentation, some of the earliest red zone pieces recovered were

LF14A and RF1 (pieces of fuselage from the CWT area, left and right sides, respectively), which were

recovered on July 24, 1996; CW504 (piece of the front spar), which was recovered on August 7, 1996; and

RF35 (piece of right-side fuselage near CWT), which was recovered on August 14, 1996. Many of the

pieces recovered from the red zone were recovered several weeks after the accident. For more information

about the green zone, see section 1.12.3. 

133 See section 1.16.3 for a discussion of the significance of crushing damage. 

134 See sections 1.12.2 and 1.12.3, respectively.

135 The floor tracks and the track support structure from the front of the forward cargo compartment were

not recovered. 

136 FQIS wiring clips (without FQIS wires) were located along the cargo compartment ceiling at the aft

end of the forward cargo compartment. 
Factual Information 70 Aircraft Accident Report

recovered cargo container stops were found intact and affixed to their respective cargo

compartment floor tracks. 

1.12.2 Wreckage Recovered from the Yellow Zone

The yellow zone was the smallest of the three zones and was contained within the

red zone on its northeastern side (see figure 22a). This zone contained pieces of the

airplaneís forward fuselage, from about STA 840 to the nose of the airplane (STA 90). The

wreckage recovered from the yellow zone included nearly all of fuselage section 41 (the

nose section) and the forward portions of fuselage section 42. Overall examination of the

fuselage pieces recovered from the yellow zone revealed severe crushing damage that

extended from the left-side main cabin floor (about 3 feet below the left-side main cabin

windows), across the bottom of the fuselage, and to just above the right-side main cabin

windows. Pieces of fuselage structure from the top of this portion of the fuselage were

found broken into pieces much larger than those from the bottom. 

Pieces of wreckage recovered from the yellow zone did not exhibit any evidence

of soot deposits or fire or heat damage. No evidence of foreign-object impact to the

cockpit windows was found. The two cargo containers that had been loaded on the right

side of the accident airplaneís forward cargo compartment were recovered from the yellow

zone. These containers were more fragmented than the two left-side containers.137

Examination of cockpit components recovered in the yellow zone revealed that the

altimeters at the captain and first officer positions (which are powered by engine-driven

generators on different wings) displayed 13,820 and 13,800 feet, respectively. The clocks

at the captain and first officer positions (which are independently set and powered by the

airplane battery in the cockpit) stopped at 2031:30 and 2031:20, respectively. 

The fuel quantity gauges at the flight engineer position indicated the following fuel

quantities: 3,100 pounds in the No. 1 reserve fuel tank; 22,200 pounds in the No. 1 main

fuel tank; 57,500 pounds in the No. 2 main fuel tank; 640 pounds in the CWT (indicated

by both the electrical and mechanical gauge mechanisms);138 60,100 pounds in the No. 3

main fuel tank; 27,300 pounds in the No. 4 main fuel tank; and 3,300 pounds in the No. 4

reserve fuel tank. The fuel-used indicators for the Nos. 1, 2, 3, and 4 engines showed

2,680; 2,710; 23,570;139 and 2,830 pounds, respectively. The fuel quantity totalizer

137 For more information, see section 1.12.1.

138 TWA flight 800 dispatch paperwork indicated that the CWT contained 300 pounds of fuel. This

reading, which would have been recorded by the ground refueler at JFK, would have been obtained from a

repeater gauge at the left wing refueling station; the repeater gauge would register the same quantity as the

cockpit gauge. The difference between the 300- and 640-pound readings exceeded the accuracy tolerance

allowed by the manufacturer and recorded during certification tests. Tests conducted by the Safety Board

during this investigation showed that the application of power to a wire leading to the fuel quantity gauge

could cause the digital display to change at a rate of about 1,000 pounds per second (direction and rate of

change varied, depending on which wires were shorted). According to circuit breaker trip curve data (see

figure 34 in section 1.16.6.2.1), more than 1,000 percent of the rated current could pass through a properly

functioning circuit breaker in 0.34 second, indicating that the digital display could change by several

hundred pounds in less time than is required to trip the circuit breaker. 
Factual Information 71 Aircraft Accident Report

indicator140 indicated a gross weight of between 587,000 and 588,000 pounds and a total

fuel-on-board weight of between 169,000 and 170,000 pounds. 

The Nos. 1, 2, and 3 crossfeed valve switches were found in the ìOnî (open)

positions, and the No. 4 crossfeed valve switch was found in the ìOffî (closed) position.141

The Nos. 1, 2, and 3 engine fuel flow indicators showed zero fuel flow, whereas the No. 4

engine fuel flow indicator showed 900 pounds per hour. 

The Safety Board examined the recovered caution and warning lightbulb

indicators142 for indications of a preaccident malfunction. Lightbulb filaments were

examined with 10 X or 15 X monocular microscope magnification or with a variable

(75 X maximum magnification) binocular microscope. These examinations revealed that

most of the lightbulb filaments, whether intact or broken, exhibited no evidence of

distortion or stretching of the filament or individual coils. 

The cabin zone temperature control panel (which is located at the flight engineerís

position) was severely damaged. The temperature control for Zone 1 was found in the

ìCoolî position (the coolest setting of the temperature controls). The trim valve indicator

for Zone 2 (see figure 6) was found at the ìMid-Heatî position (the warmest setting of the

temperature controls). The instrument panel was broken and bent where the air

conditioning pack selector switches should have been attached, and the pack selector

switches were attached by their respective wire bundles. All three switches were bent

upward toward the 12 oíclock position. 

Although postaccident examination revealed that the cockpit flap control lever was

located between the 10° and 20° extended positions, physical damage to recovered leading

and trailing edge flap components indicated that the flaps were in the retracted positions at

the time of water impact. The landing gear handle in the cockpit was in the center position,

marked ìOff,î and physical damage indicated that the landing gear was retracted at the

time of water impact. The pitch-trim jackscrew was found about 3/4 up from the bottom of

the jackscrew (10 screw threads visible above the carriage, 35 threads visible below),

which, according to Boeing, is consistent with a climb-trim position. 

1.12.3 Wreckage Recovered from the Green Zone

The green zone was located farthest east (farthest from JFK) in the wreckage

distribution. Most of the airplane wreckage was recovered from this zone, including most

of the pieces of both wings, all four engines, and the fuselage aft of about STA 1000

139 The fuel-used indicator for the No. 3 engine was heavily damaged. 

140 As previously indicated, the TWA 747 Operations Manual describes this flight engineer station

indicator as a gross weight/total fuel weight indicator. 

141 Although individual switch positions throughout the cockpit wreckage were documented, most of the

switches were not locked in position, and many had strands of wire wrapped around the switch toggle. 

142 Many of the caution and warning lightbulb indicators were not recovered, and some of those that were

recovered were badly damaged. Loss of power can also affect filament analysis. See Systems Group

Chairmanís Factual Report, dated November 17, 1997, for detailed information. 
Factual Information 72 Aircraft Accident Report

(fuselage sections 44, 46, and 48ósee figures 3a and 3b for reference). In addition, some

pieces from the aft portion of fuselage section 42 were recovered from the green zone. The

green zone also contained most of the WCS/CWT, including all recovered pieces from the

rear spar, SWB1, the mid spar, SWB2 (except the SWB2 manufacturing access door,

which was recovered from the red zone), and the upper and lower skin panels of the WCS.

Portions of SWB3 and the front spar were also recovered from the green zone, as were the

aft portion of the keel beam, the aft air conditioning pack, the main landing gear, and the

tail section. 

Examination revealed that some pieces of wreckage recovered from the green zone

exhibited heavy soot deposits and severe heat damage. Fire and heat damage (including

blackened structure, melted wiring and aluminum materials, and burned composite

materials) and heavy soot deposits were found on the following: some pieces from the aft

CWT; floor beams and some of the passenger seats located just aft of the CWT rear spar;

the portion of the fuselage above the right wing root, including the airplaneís external

skin, parts of the right wing (including the front spar), and portions of the left wing

outboard of the No. 1 engine.143 Appendix F contains soot deposit/fracture diagrams for

the WCS, wings, and fuselage. 

The fire damage was particularly severe on a portion of the right-side fuselage

structure that remained attached to the inboard section of the right wing (including the

fuselage area surrounding the R3 cabin door, located just above the right wing root).

Heavy soot deposits were found on the top surface of the right wing. The R3 cabin door,

which had separated from the fuselage and was recovered separately, exhibited fire and

heat damage and soot deposits on the inside and outside surfaces with some areas of

melted-through aluminum. Portions of the door normally shielded by the door frame and

portions of the door frame normally shielded by the door also exhibited heavy soot

deposits. In addition, severe fire damage was observed on the left side of SWB2, large

portions of which appeared to have been melted away. 

Examination of the WCS upper skin panel pieces revealed that pieces from the

forward left side of the panel exhibited no soot deposits, whereas pieces from the right

side of the upper skin panel exhibited moderate to heavy soot deposits. Inspection of

pieces of the WCS lower skin panel and the aft portion of the keel beam revealed

substantial deformation and some soot deposits. Specifically, most of the lower surface of

the WCS lower skin panel exhibited widespread moderate to heavy soot deposits, with the

soot accumulation generally heavier on the right side; most of the upper surface of the

lower skin panel exhibited light soot deposits, with localized areas of heavy soot deposits. 

Most right and left wing wreckage pieces, including the wing tip antennas, were

recovered from the green zone, although some small internal pieces of the left wing were

recovered from the red zone during trawling operations (late in the recovery process), and

some pieces of the right and left wings, including the outboard portion of the right wing,

were found floating (early in the recovery process) and were not associated with a specific

143 The fire and heat damage observed outboard of the No. 1 engine on the left wing is discussed further

later in this section. 
Factual Information 73 Aircraft Accident Report

debris zone. The wing landing gear and engines were not attached to the wings when they

were recovered. Most pieces of the right and left wing flight control surfaces (including

leading and trailing edge flaps, ailerons, and spoilers) were recovered from the green

zone144 and exhibited general impact damage. Many portions of the trailing edge flaps,

ailerons, and spoilers and some portions of the leading edge flaps exhibited fire damage

and soot deposits. 

The outboard sections of both wings had separated from their respective wings just

outside of the outboard engines, near wing station 1224. Both separated outboard wing

sections were recovered relatively intact; they were about 29 feet long and extended from

about wing station (WS) 1224 to the wing tip. 

Examination of pieces of the right wing revealed the presence of soot deposits in

the fuel tank vent stringers in the right wing structure, the separated outboard portion of

the wing, and in the surge tank in the wing tip. Heavier soot deposits were observed on the

outboard portions of the attached wing structure, while the separated outboard portion of

the right wing and the vent stringers closer to the right wing root contained light soot

deposits. Additionally, fire (heat) damage was observed on the upper skin of the right

wing section outboard of the No. 4 engine to the location where the most-outboard portion

of the wing separated. No fire (heat) damage was observed on the separated outboard

portion of the right wing. 

Most of the left wing stringers (including vent stringers) between the left side-ofbody rib and WS 1224 had separated from the upper and lower skin pieces; these stringers

were bent and curled in various directions. The left wing outboard of the No. 1 engine was

fragmented and was recovered in many pieces. The most-outboard piece of the left wing

that was recovered was a piece of the left wing tip that included the top and bottom wing

surface near the surge tank. Light soot deposits were observed on the interior of this piece

of wreckage, near the front spar and on the lower surface of the surge tank, near the vent

stringer. The fracture surfaces in these areas did not contain soot deposits. Additionally, a

relatively large piece of the lower left wing surface from just outboard of the No. 1 engine

was recovered; although the internal and fracture surfaces on this piece were clean, light

soot deposits were observed on its external surface. The lower surface of the left wing

contained soot deposits, with moderate to heavy soot deposits just inboard of the bottom

fairing. The upper surface of the left wing was mostly free of soot. The upper skin of the

inboard section of the left wing was fractured into numerous pieces, whereas most of the

lower skin on the inboard portion of the left wing was recovered in larger pieces. 

Fuselage section 44 is located above the wings and extends from the front spar to

the aft end of the wheel wells (STA 1480). The 16 main landing gear tires were recovered

from the green zone and examined at the hangar in Calverton. All of the tires were found

burst with evidence of external forces and exhibited cuts and gouges that were consistent

with the recovered wheel well wreckage. No evidence of internally generated failures

associated with heat was found on any of the tires.145

144 Some pieces of flight control surfaces were recovered floating on the oceanís surface and were not

associated with a specific debris zone. 
Factual Information 74 Aircraft Accident Report

All seven cargo containers that had been loaded in the accident airplaneís aft

container cargo compartment were recovered from the green zone and were severely

damaged in a manner consistent with impact forces. Six of these seven cargo containers

were constructed of aluminum and exhibited a similar degree and nature of damage. The

seventh cargo container was constructed of fiberglass-reinforced plastic, instead of

aluminum, and was shattered. 

Fuselage section 46 extends between STA 1480 and 2360. The recovered pieces of

the fuselage from this section exhibited a fragmentation pattern similar to that observed on

the forward portion of the fuselage, which was recovered in the yellow zone. The upper

fuselage structure in this section was found broken into relatively large pieces, while the

lower fuselage structure (below the window belt on the right side of the fuselage and

below stringers 28 to 31 on the left side of the fuselage) was fragmented into smaller

pieces. Most upper fuselage skin pieces were totally or partially separated from the frames

and stringers. The lower fuselage pieces exhibited inboard bulging in the skin bays (the

area between adjacent stringers and frames), with stringers broken or damaged at most

frame stations. 

Fuselage section 48 broke into large sections, the majority of which did not exhibit

substantial compression damage. Minimal evidence of soot deposits or heat damage was

found on pieces of wreckage from the aft portion of sections 46 and 48. 

1.12.4 Engines 

The four engines were found in the green zone separated from the wings. The

No. 1 engine had crushing damage along the bottom of the nacelle and the low-pressure

compressor (LPC), high-pressure compressor (HPC), and diffuser cases; the No. 2 engine

had crushing damage along the right side of the nacelle and the LPC, HPC, and diffuser

cases; the No. 3 engine inlet and fan blades were crushed rearward, and the LPC and HPC

cases had crushing damage from the front and along the bottom; and the No. 4 engine inlet

was crushed rearward, the LPC and HPC cases were crushed axially from the front and

inward from the right side, and the diffuser case was crushed along the right side. All of

the fan cases and cowls were separated from their respective engines. All engine thrust

reversers were also separated from their respective engines; however, examination of the

recovered thrust reverser actuators showed that the drive mechanisms were at the head end

of the jackscrew, consistent with a thrust reverser stowed position at the time of impact. 

The Safety Boardís disassembly and examination of the four engines revealed LPC

damage consistent with a minimal amount of low-pressure rotor rotation (if any) at the

time of impact and HPC damage consistent with some high-pressure rotor rotation at the

time of impact. No evidence of uncontainment, case rupture, fire, penetration of an object

from outside into the engine, or preimpact damage was found in any of the engines.

145 Although there was no evidence of heat-related damage, investigators observed marks on the surfaces

of the four tires mounted on the left wing main landing gear that were not seen on the other tires. (In their

stowed positions, these tires would have been located just aft of the CWT.) These marks appeared similar to

light soot deposits on the rubber material and exhibited no evidence of exposure to sustained heat or fire. 
Factual Information 75 Aircraft Accident Report

Additionally, the Boardís examination of the APU revealed no evidence of uncontainment,

case rupture, fire, penetration of an object from outside into the APU, or preimpact

damage.146

1.12.5 Center Wing Fuel Tank Pumps and Other Components 

Both jettison/override pumps from the CWT were recovered, examined, and

tested. The jettison/override pumps were impact damaged (especially at the aft ends,

where they had been attached to the rear spar), but exhibited no evidence of electrical

shorts, arcing, preimpact breaching of the housing, or discoloration of the pump shaft.

(Figure 23 is a diagram of a jettison/override fuel pump.) Examination revealed that the

jettison/override pumpsí thermal fuses and flame arrestor tubes were intact and exhibited

no evidence of thermal or flame exposure. The CWT jettison/override pump motors

exhibited no evidence of heat-related discoloration; the motor windings were the same tan

color as those from the wing fuel tanks. (A similar jettison/override pump was heated in a

laboratory test until the thermal fuse melted at slightly less than 400° F, and the motor

windings were subsequently found to have darkened substantially.) 

The CWT jettison/override pump control switches in the recovered cockpit

wreckage were found in the ìOffî position. Investigators examined the possibility that any

one or all of these switch positions might have been altered during the accident sequence.

Unlike the scavenge pump switch (which was recovered damaged in the ìOffî position),

the jettison/override switches did not have a locking feature, and the recovered switch

positions could not be positively related to a preimpact position. The jettison/override

switch base damage was consistent with the toggle switch being in the position to which it

was spring-loaded and not at the ìOffî position. However, according to TWA procedures,

with only unusable fuel remaining in the CWT, these pump switches should have been off

at the time of the accident. Further, the CVR did not record any comments regarding

activation of these pumps.

Although the general condition of the left CWT jettison/override pump was clean,

two pieces of a clear, soft, rubberlike foreign material were found on the outside diameter

of the motor/impeller housing. Infrared examination by the National Aeronautics and

Space Administration (NASA) determined that the material was a silicone-based

substance. Additionally, the right jettison/override pump exhibited evidence of rubbing

between the impeller edges and adjacent housing surfaces. Overhaul records from Crane

Company, Hydro-Aire, Inc. (the pump manufacturer), described similar damage in other

pumps and attributed it to worn shaft bearings.147

146 For additional information, see Powerplants Group Chairmanís Factual Report, dated February 25,

1997. 

147 Similar, but extensively deeper rubbing damage was found in jettison/override pumps in the

Philippine Airlines 737 that had a CWT explosion on May 11, 1990. The Safety Board conducted tests for

2 weeks to determine whether galling between aluminum surfaces within the pump could have ignited

flammable vapors. The investigative group was not able to ignite flammable vapors by impeller rubbing or

by other types of contact between internal fuel pump components. 
Figure 23. A diagram of a jettison/override fuel pump. FORWARD
Factual Information 77 Aircraft Accident Report

The CWT scavenge pump motor and impeller were not recovered. However,

examination of the recovered fragments of related components, including the

flange/spacer that was mounted between the scavenge pump and the rear spar, did not

reveal evidence that the scavenge pump could have been an ignition source. Examination

of the rear spar where the flange would have been mounted revealed a flange-shaped clean

area, surrounded by an area of heavy sooting, with evidence of heat damage. There were

no soot or fire patterns emanating from the scavenge pump mounting hole in the rear spar,

and there was no indication that the heat detectors in the main landing gear wheel well had

detected heat. 

A 10 3/4-inch-long section of fuel tube assembly with a P/N unique to the CWT

scavenge pump pick-up tube was recovered. According to an installation drawing, the

tube assembly would have been installed near an opening in the mid spar. Although the

outside of the tube contained soot deposits, the internal surfaces had a slightly golden

color, similar to the anodize finish seen on other fuel tubing. No evidence of soot deposits

or flame flow patterns were observed in the tube that could be related to an ignition

source. 

1.12.6 Air Conditioning Equipment 

Air conditioning packs 1 and 3 (the forward-most air conditioning packs) were

recovered from the red zone, and pack 2 was recovered from the green zone. Examination

of recovered pieces of the three conditioned air systems revealed that numerous ducts and

the heat exchanger headers exhibited flattening on the upper surfaces. However, there was

no evidence of preimpact leaks that may have abnormally heated the CWT or of an

uncontained rotor burst from the ACMs. Disassembly of the ACMs revealed no evidence

of preimpact rotational scoring in housing areas adjacent to the impellers. Where the

ducting that routed conditioned air from the air conditioning packs to the plenum was

found broken or ripped, no localized color changes were observed on the ducting or

adjacent materials. Where breaks occurred in heat-affected weld zones on this ducting, the

Safety Boardís metallurgists found no evidence of fatigue failures. 

Some of the titanium ducts that transport hot bleed air from the engine(s) or APU

to the air conditioning packs were examined in the Safety Boardís Materials Laboratory

and found to contain preexisting cracks. According to Boeing, the blue coloration on the

exterior surface of the ducts and on the preexisting crack areas indicated that the fractures

existed when the ducts were stress relieved during manufacture. On some of the duct

fractures, the preexisting cracks extended all the way through the duct wall thickness. The

circumferential length of these cracks ranged from between 0.1 to 0.5 inch. There was no

evidence that the cracks that extended through the duct wall had extended around the

circumference of a duct before the accident. 

The turbine bypass valves for the three air conditioning packs were positioned

about 10° from the fully closed position. The two recovered ram air inlet doors were found

in the fully open position, and the two recovered ram exit actuators were found in the fully

retracted (door open) position. According to the Hamilton Standard Service Manual
Factual Information 78 Aircraft Accident Report

(HSSM), these positions indicated that the three air conditioning packs were operating to

cool the airplane cabin. The recovered flow control valves were found in the closed

position. According to the HSSM, a spring would move the flow control valve to the

closed position if it did not have positive pneumatic pressure. 

In the 747-100, some conditioned air is moved forward from the air duct/plenum

chamber (which is located above the WCS and beneath the cabin floor) through what

investigators referred to as ìthe air conditioning long duct,î which begins about STA 1000

(above the front spar) and continues to about STA 1080 (between SWB3 and SWB2). The

forward portion of the air conditioning long duct had broken, and three pieces (totaling

about 67 inches in length) were recovered and examined. Most of the polyurethane

insulation that normally surrounds the air conditioning long duct was missing from these

pieces. 

1.12.6.1 Brown Splatter Material on Air Conditioning Ducts

The forward-most piece of the air conditioning long duct was about 14 inches long

and exhibited little evidence of fire or heat damage. There was brown splatter material on

the outside of the left and lower duct surfaces and on the inside lower surface, extending

3 inches forward from the aft end of this piece of duct. The splatter material continued

across the aft fracture surface. The middle piece was about 28 inches long (with the aft end

extending to above SWB3) and contained some soot deposits. Brown splatter material

extended aft from the forward end of this piece as follows: about 6 inches aft on the

right-side inside duct surface, about 12 inches aft on the inside lower surface, and about

16 inches aft (intermittently) on the left-side inside duct surface. The aft-most piece was

about 25 inches long and contained darker soot deposits than the middle piece. Brown

splatter material extended aft about 4 inches from the forward end on the inside surface of

this piece. Some brown splatter material was observed on the inside upper and lower

surfaces at the aft end of this piece of duct; the splatter material on the inside upper surface

extended forward 8 to 10 inches in spots. 

At its forward end (above the front spar), the air conditioning long duct connects to

what investigators called the S duct, which curves upward and outboard from the long

duct. The aft portion of the S duct, extending 21 inches forward of the air conditioning

long duct, was recovered and examined. The polyurethane insulation surrounding the

recovered portion of the S duct was damaged and, in some places, missing. In areas where

the insulation was missing, the duct appeared to have been scraped. The forward end of

the recovered S duct piece was split longitudinally, and the inside surface of the duct

exhibited numerous brown splatter material. Similar brown splatter material was also

observed on the air conditioning long duct, the left side of the forward upper WCS and the

front spar, and on some cabin passenger seats located above and/or forward of the WCS. 

The brown splatter material on the upper WCS extended aft along the air

conditioning long duct to the point where it met the trim air ìYî inlet. At that location, the

insulated metal trim air pipe is attached to the air conditioning long duct by a connector.

Examination of the recovered trim air pipe revealed that a portion of the connector was
Factual Information 79 Aircraft Accident Report

still attached to the end of the pipe, and no evidence of an overheated condition was

found.148

1.12.7 Recovered Electrical Components/Wiring 

The Safety Board thoroughly examined all recovered and identified electrical and

FQIS components and wiring (including connections, repairs, bundling, bundle fragments,

various loose wires).149 Almost every piece of recovered wire exhibited damage, including

numerous cracks in the insulation (many of which penetrated to the core conductors),

extensive abrasions, twists, and cuts; further, salt deposits were found on insulation and on

copper where insulation had been breached. (Figure 24 shows wreckage recovered with

tangled and damaged wires attached.) Some of the recovered wiring exhibited evidence of

electrical arcing; however, there was no evidence of electrical arcing to recovered CWT

FQIS wiring (about half of the CWT FQIS wiring was recovered and identified).

Microscopic examination (up to 70 X) revealed that the ends of most wires were necked or

sheared. Blackened insulation was observed on many wires; however, many of these wires

had been routed through the burned areas of the fuselage and right wing root. Figure 18 in

section 1.6.1.3 is a general illustration of 747 wire raceways between the cockpit and the

aft fuselage, taken from TWAís Electrical/Electronic Wiring Raceway Identification

[91-00-00]. 

Most of the wiring recovered from the accident airplane was found separated from

its mounting structure. Some wire fragments in the wings remained attached to the wing

structure by production clips, and other wires got tangled with structural wreckage. The

majority of the forward fuselage wiring was received at the hangar in Calverton in two

sections. Both of these sections were tangled with other debris, which included structural

materials and other electrical components; one section included material from the cockpit

and upper deck, whereas the other section included material mostly from the E/E bay that

had been located between the nose landing gear wheel well and forward cargo

compartment.

Because the right side (the flight engineerís side) of the accident airplaneís forward

fuselage (including the wiring and electrical/FQIS components in that area) was severely

fragmented, recovery and documentation of all of the electrical and FQIS components and

airplane wiring from that area were not possible. Recovered electrical components from

the flight engineer station included the CWT fuel panel (including the quantity gauge and

fuel flow indicators), the FQIS wiring connector at the flight engineer station, the resistor

installed between the FQIS wire and the attached AIDS wiring,150 and most circuit

breakers (most of which were broken). 

148 For additional information regarding the brown splatter material, see section 1.16.4.6.

149 For additional information regarding wiring in transport-category airplanes, see section 1.18.2. 

150 Resistors are used in circuits to limit the amount of current (amperage), and thus power, that can pass

through them. According to Boeing, protective resistors connected to 747 CWT FQIS wiring (including the

fuel quantity gauges, VSO unit, AIDS, and fuel quantity totalizer gauge) limit the amperage from an internal

short circuit to about 1 milliamp.
Factual Information 80 Aircraft Accident Report

Some wiring from the upper left fuselage aft side of the cockpit was recovered in

loose fragments, whereas other wiring from this area was recovered tangled with

structural debris. The upper deck electronics rack (E11) was recovered as an independent

piece of wreckage, with fragments of wire bundles attached. No evidence of fire damage

or soot deposits was observed on the recovered forward fuselage electrical components or

wiring. 

The BMS13-42A Poly-X wire from the accident airplane was stiff and inflexible,

whereas recovered Teflon wires (used in the accident airplaneís fuel tanks) and samples of

BMS13-42A Poly-X wire obtained from a retired 747 (N93117) were supple and pliable.

Recovered wiring from the area near the CWT was tangled and exhibited fire damage and

soot deposits. About half of the original FQIS wires (as shown in Boeingís 747-100

engineering and production records) were recovered and identified. Although no evidence

of arcing was observed on any of the recovered components connected to the CWT FQIS

(the fuel quantity indicator, totalizer gauge, AIDS, VSO, and left wing refueling station),

examination of other FQIS wires and wires routed adjacent to the FQIS wires revealed

possible evidence of arcing in several locations. For example, examination of the

recovered wiring and wreckage from the right wing revealed evidence of arcing on

Figure 24. A photograph showing wreckage recovered with tangled and damaged 

wires attached.
Factual Information 81 Aircraft Accident Report

generator cables routed with a wire bundle in the leading edge of the right wing, near the

wing root. The right wing front spar was fractured and exhibited evidence of heat damage

in the area. The wire bundles had melded into a mass that included wires leading to the

right main wing tank (No. 4) fuel flow gauge and also included right wing FQIS wiring

that would have been routed through the right wing fuel tanks to a common connection

with the CWT FQIS at terminal strip T347. 

In addition, two non-FQIS wires that would have been corouted in the same

raceway with the CWT FQIS were found with possible arcing or heat damage at

STA 955.151 Copper strands on these two wires had melted and rehardened. The wire

bundle fragment had fire damage. None of the adjacent wires in this section of the

raceway, including some recovered FQIS wiring, exhibited the same

melting/resolidification; however, some FQIS wiring from this area was not recovered.152

Fragments from all seven of the CWT fuel quantity probes were recovered; two of

these fuel quantity probe fragments contained pieces of their respective terminal blocks.

Examination of the recovered terminal block fragments revealed that they exhibited

characteristics of the Honeywell Series 4 and subsequent terminal blocks (for example,

smooth [unknurled] surface areas and a nylon strain relief clamp instead of a metal

clamp).153

An intact compensator was recovered still attached to fuel tank structure from the

No. 4 reserve fuel tank; the compensator and fuel tank structure exhibited evidence of fire

damage. In addition, a large fragment of another compensator, exhibiting evidence of heat

and fire damage, was recovered. This compensator fragment had no identification

markings that would associate it with a specific fuel tank; however, the only pieces of fuel

tank structure near compensator installations that exhibited evidence of fire damage were

from the No. 4 reserve fuel tank and the CWT. Further examination of this compensator

fragment revealed that electrical parts inside the attached support tube had soot deposits

and exhibited evidence of melting and resolidification and blackening at the bottom edges

151 Maintenance records revealed that several repairs had been conducted in this area, including

structural repairs from a burst potable water tank and numerous other floor repairs. Metal shavings were

found adhered to recovered fragments of a floor beam from STA 920, within 2 inches of where the CWT

FQIS wiring would have been routed in the raceway. 

152 Another location in which evidence of possible arcing might have been observed involved three left

wing FQIS wires from a bundle near STA 510. Investigators located a note, written by a member of the

Systems Group, which indicated that he observed evidence of ìarching (sic)î on the FQIS wires. (The

747-100 CWT FQIS is not isolated from left wing FQIS wiring; during postaccident evaluation of wiring

paths on a retired 747, an electrical signal placed on the CWT FQIS wiring was also detected on wiring from

the left wing FQIS wiring.) However, subsequent attempts to locate the relevant portion of this wire bundle

to confirm this observation were not successful, and some Systems Group members did not recall seeing any

such evidence of arcing. (Some other Systems Group members did recall seeing evidence of arcing.) Some

parties to the investigation did not believe that this note should be considered significant evidence of arcing. 

153 Honeywell Series 1 to 3 fuel quantity probe terminal blocks had knurled surfaces (relatively sharp

pointed cones in the hard plastic) that resulted in sharp edges resembling saw teeth at the terminal block

edges. Honeywell Series 4 (and subsequent series) fuel quantity probe terminal blocks did not have these

knurled surfaces. Additionally, wires were attached to Series 1 to 3 fuel quantity probe terminal blocks with

metal clamps, whereas wires were attached to Series 4 fuel quantity probe terminal blocks with nylon-lined

clamps. 
Factual Information 82 Aircraft Accident Report

of the plastic parts. The soot deposits exhibited an upward flow pattern and were located at

the lowest edges of a plastic sleeve that covers a wire shielding connection, two plastic

insulators on connectors, and the plastic lead wire support. The general features of each

plastic part did not appear to have been molten, and away from the blackened edges each

part retained its color. (Recovery records indicated that this intact compensator and the

compensator fragment were recovered from the green zone.) 

Investigators found that Boeing and TWA wiring documents for the accident

airplane did not show all of the wires that were recovered.154 Additionally, examination of

wire configurations in the accident airplane revealed that they did not match those shown

in Boeingís technical diagrams. For example, Boeingís PI 61B70126 showed separated

wire bundles in a tray located above the CWT. Portions of these bundles recovered from

the accident airplaneís wreckage were found twisted in a single mass of wires, with no

evidence that they had been separated into wire bundles as illustrated in Boeingís PI.

Further, the Safety Boardís examination of this area of wiring in other 747s (N93117 and

N93105), including some operated by TWA, revealed that the wires were apparently not

bundled as shown in the PI but were randomly laid in trays.155

Examination revealed black spots on crimped FQIS wire connectors, including one

on a damaged fuel probe from the accident airplaneís CWT. (During the Safety Boardís

participation in the investigation of a May 1976 Iranian Air Force 747 crash near Madrid,

Spain, investigators observed a similar black spot on a wire located adjacent to the site of

this crimped wire connector.)156 X-ray energy dispersive spectroscopy (EDS) examination

of the black spot revealed that it contained copper, sulfur, and trace amounts of silver.157

According to fuel experts at the U.S. Air Force (USAF) Research Laboratory (AFRL) at

Wright Patterson Air Force Base, this residue was similar to silver-sulfide deposits158 that

they observed during previous examinations of fuel probes from military airplanes. 

The Safety Boardís examination of wiring from the accident airplane revealed

evidence of damage to electrical insulation on wires that had been attached to FQIS fuel

quantity probe terminal blocks. Sulfide deposits were found on some exposed conductors.

Much of the damage, which included exposed conductors, was not visible until the wiring

was removed from the fuel quantity probe terminal blocks for inspection.159

154 Investigators recovered unmarked wiring, which was not shown on the Boeing or TWA documents. 

155 For additional information about the Safety Boardís inspections of other airplanes, see section 1.18.2.

156 See National Transportation Safety Board. 1978. Special Investigation Report ñ Wing Failure of

Boeing 747-131, Near Madrid, Spain, May 9, 1976. NTSB/AAR-78/12. Washington, DC. 

157 For the purposes of this report, deposits containing copper, sulfur, and silver will be called

silver-sulfide deposits. Silver-plated copper wiring is used in the 747 FQIS. Sulfur is commonly contained in

jet fuel. In April 1998, Boeing provided the Safety Board with laboratory reports and other documents that

described the presence of sulfur in turbine fuel as early as September 1970 (Budd, B. J. and Sanger, R. P.

Silver Corrosion by Aviation Turbine Fuel. Institute of Petroleum.).

158 For additional information regarding sulfide deposits, see section 1.16.6.9.

159 Similar damage was observed on several other 747 airplanes during postaccident inspections. Some

of the exposed core conductors apparently resulted from cold flow (migration or displacement) of the Teflon

insulation material, whereas others were apparently the result of mechanical damage. 
Factual Information 83 Aircraft Accident Report

Additionally, investigators noted that various methods had been used to route

wires to and secure them at the fuel quantity probe terminal blocks. In general, two

different types of fuel quantity probe terminal blocks were recovered in the accident

airplane wreckageóthose with knurled surfaces and those without knurled surfaces. (See

figure 15.) Investigators observed instances in which adjacent FQIS insulation had been

breached near the knurled area of some terminal blocks. Honeywell documents (the

Component Maintenance Manual and Overhaul Manual) indicated that both terminal

block designs were in service at the time that the accident airplane was manufactured.

According to an October 27, 1997, letter from Boeing, ì[a] production change was made

at Boeing that installed the Series 4 probes in line number 65 and on. [The accident

airplane] was line number 164 and was delivered 10/27/71, so it is improbable that it was

delivered with Series 3 terminal block probes.î 

Examination of the recovered FQIS wires revealed wires that had been repaired for

damage that existed before the accident, evidenced in some cases by splices of various

types. Several wire repairs were also found that did not comply with standards for repairs

or installations used by Boeing or TWA, including the following:160

ï An oversized strain relief clamp was used on the terminal block of the

compensator in the No. 1 fuel tank. The cable harness passing through the

clamp was wrapped around to pass through the clamp twice and still was not

firmly secured. Chafing was observed on some wires. 

ï Numerous wire splices were covered by a plastic insulating sleeve over metal

barrels, such that the ends of the wire splices were open, with no other

wrapping. These wire splices were noted at locations throughout the airplane

(including areas exposed to fluid contamination, such as above the potable

water tanks on the front spar) and occasionally showed evidence of corrosion. 

ï Several wire bundles contained numerous wire splices on adjacent wires at the

same location. Boeingís SWPM recommended that wire splices be separated

by a minimum of 0.25 inch, and Boeing SL 747-SL-20-048, dated January 25,

1995, stated that ì[wire s]plices should be staggered throughout the bundle,

this reduces the potential for wire-to-wire chafing and bundle interference.î

ï The connector pins in the recovered fuel totalizer gauge contained excessive

solder, which appeared to have inadvertently joined connecting pins/wires

from the right wing main fuel tank and CWT FQIS. The excessive solder had

cracked between the connecting pins.161

160 During accident-related inspections of other transport-category airplanes, the Safety Board observed

conditions and repairs similar to these in a variety of other airplanes operated by different airlines. For

additional information, see section 1.18.2. Also see Systems Group Chairmanís Factual Report, dated

November 17, 1997; Systems Group Chairmanís Factual ReportóAddendum for Aircraft Wire Inspections

and Historical Reports, dated July 28, 1999; and System Group Chairmanís Factual ReportóAddendum for

Fuel Quantity Indicators, dated February 5, 2000. 

161 Although the cause of the cracking in the solder is unknown, postaccident tests conducted at

Honeywell (the manufacturer of the fuel totalizer gauge) revealed that when slightly more than 270 volts

was applied to one of the connecting pins, electrical energy would cross the crack in the solder. 
Factual Information 84 Aircraft Accident Report

1.13 Medical and Pathological Information

The bodies of 99 of the 230 airplane occupants were recovered from the surface of

the ocean by various civilian, military, and police vessels during the first 24 hours after the

accident. U.S. Navy and local police divers discovered most of the remaining victims

during the next 3 months; however, some remains were recovered during the trawling

operations (the final stage of search and recovery operations, which continued until

April 30, 1997). The last recovered human remains were retrieved by a fishing trawler on

May 22, 1997. In most cases, when victims were recovered, they were transferred from

the dive boats and/or salvage ships to a Suffolk County Police Department boat and

transported to the Suffolk County Medical Examinerís Office in Hauppauge, New York.162

In some cases, victims were transported to shore by helicopter. 

Most identifications of occupants were accomplished through the use of

fingerprints or dental records. However, in 29 cases, neither of these methods was

successful; these cases required the use of deoxyribonucleic acid (DNA) protocols or

forensic radiography as the primary means of identification. (Nineteen occupants were

identified solely by DNA, and 10 were identified by forensic radiography, either by the

medical examiner or the Armed Forces Institute of Pathology.) 

Toxicological samples (muscle tissue) from the captain (left front seat), the flight

engineer trainee, and the check flight engineer were sent to the FAAís Civil Aeromedical

Institute (CAMI) in Oklahoma City, Oklahoma, for examination. Suitable toxicological

specimens were not available for the captain/check airman (right front seat). The

toxicological results for all submitted specimens were negative for all drugs of abuse163

and for prescription and over-the-counter medications. CAMIís toxicological report

indicated that the presence of small amounts of alcohol in some of the specimens was

most likely ìfrom postmortem ethanol productionî caused by decomposition. 

During the investigation of the accident, the Medical Forensic Group, which

included medical/forensic experts from the USAF, the U.S. Navy, the U.S. Army (retired),

the FBI, the Safety Board, and the Suffolk County Police Department, was formed to

review the available medical and forensic documentation for evidence that an explosive

device (bomb or missile) had detonated near any passenger or crewmember. The review

revealed no localized areas of damage or injuries in the airplane.164 In addition,

investigators used biomechanical analysis and correlated medical and forensic data with

passenger seat assignments165 and recovered seat damage (422 of the 455 seats on the

162 A temporary morgue at the U.S. Coast Guard (USCG) station at East Moriches was used during the

first several weeks of the investigation. The remains of the occupants of TWA flight 800 that were recovered

subsequently were transported to the Suffolk County Medical Examinerís Office in Hauppauge, New York,

for identification and examination. 

163 The five drugs of abuse tested in postaccident analysis are marijuana, cocaine, opiates, phensyslidine,

and amphetamines. 

164 If an explosive device had detonated inside the airplane, localized areas of damage/injuries would be

expected.
Factual Information 85 Aircraft Accident Report

airplane were recovered) to reconstruct injury events as they occurred during the

explosion, breakup, and water impact of the accident airplane. 

A Medical Forensic Investigation Analysis Report, dated January 28, 1999, and

prepared for the Department of Justice/FBI by a medical/forensic expert,166 concluded the

following: 

ï Exhaustive analysis of all available medical data on the victims of TWA Flight

800 by an experienced team of forensic pathologists, biomechanicists and

criminal investigators failed to find any evidence that any victim was directly

exposed to a bomb blast or missile warhead detonation. This finding makes it

highly unlikely that a localized explosion occurred within the passenger cabin

of TWA Flight 800.

ï All injuries found in the victims were consistent with severe in-flight break up

and subsequent water impact.

ï Injury and burn patterns to the victims as well as some body locations suggest

that there was a severe break up of the passenger cabin early in the crash

sequence.

ï Fire propagated in the [cabin zone located above the CWT] after most

occupants [of this section] had been ejected. The small number of passengers

with burn injuries exhibited only superficial burns consistent with exposure to

a flash flamefront. 

1.14 Fire/Explosion

Physical evidence indicated that an overpressure event167 occurred in the airplaneís

CWT.168 During the resultant airplane breakup, additional fires occurred. Witness

documents169 and physical evidence also indicated that, after the wreckage impacted the

water, fuel on the waterís surface continued to burn for several hours. 

Subsequent examination of the recovered airplane wreckage revealed that

significant fire damage was observed only in specific areas, including components from

within the CWT, floor beams and some of the passenger seats that had been located above

and just aft of the WCS, the portion of the fuselage adjacent to and above the right wing

165 The Safety Board notes that passenger seat assignments may not reflect the seats actually used by all

passengers. The accident flight was not a full flight, and passengers may have moved during the ground

delay at JFK. 

166 Shanahan, Dennis F., M.D., M.P.H., Colonel, U.S. Army, Medical Corps (retired). For additional

information, see Dr. Shanahanís complete report and the Medical Forensic Group Chairmanís Factual

Report, dated October 17, 1997. 

167 In this report, the Safety Board uses the term ìoverpressure eventî to connote an event in which the

pressure in the CWT is increased in a relatively short time to a level at which the structural integrity of the

CWT was compromised. 

168 For breakup sequence information, see section 1.16.3.2 

169 For additional information regarding witnesses and witness documents, see section 1.18.4. 
Factual Information 86 Aircraft Accident Report

root, portions of the right wing, and portions of the left wing outboard of the No. 1 engine.

Soot deposits on the fuselage aft of the front spar were only observed on the external

fuselage skin; however, heavy soot deposits were observed on the aft portion of the keel

beam.170

Additional fire and explosion information is contained in other sections of this

report, as pertinent. For example, section 1.12 contains descriptions of the fire, heat, and

soot damage observed on recovered wreckage. Additionally, sections 1.16.1, 1.16.2, and

1.16.3 contain radar, trajectory study, and sequencing study information, respectively,

regarding the airplaneís breakup. Also, sections 1.16.5 and 1.16.6 contain additional

information regarding heat- and fire-related damage/evidence, research into the fuel/air

mixture on board the accident airplane at the time of the accident, the characteristics and

combustion behaviors of Jet A vapors,171 and research and testing to identify the potential

ignition location. See appendix F for documentation of soot patterns on pieces of

wreckage and appendix G for a list of other fire/explosion-related airplane accidents. 

1.15 Survival Aspects

Although individuals in various civilian, military, and police vessels reached the

accident site and initiated a search for survivors within minutes of the initial water impact,

no survivors were located. The accident was not survivable.172

170 For more detailed information regarding fire and explosion damage, see Fire and Explosion Group

Chairmanís Factual Report, dated August 22, 1997. Also, see Fire and Explosion Group Chairmanís Factual

ReportóAppendixes I and II (Sooting/Fracture Diagrams). 

171 Jet A fuel is a kerosene fuel used in civilian turbine engine airplanes. It is composed of a mixture of

more than 100 distinct types of hydrocarbon molecules; the precise composition often varies between

refinery and by season. Jet A fuel is specified to have a minimum flash point of 100o F. Jet A-1 is a similar

fuel, but has a slightly lower freezing point. Although Jet A fuel is available in some other countries, it is

used primarily in the United States.

172 Although some of the accident airplane occupants may not have received injuries that were

instantaneously fatal, all occupants received fatal injuries; therefore, the accident was considered

nonsurvivable. Sufficient remains were recovered of 202 of the 230 occupants for investigators to make a

determination of the level of severity of the injuries received. The following severity criteria was used:

severe trauma was defined as ìtrauma that resulted in instantaneously fatal injuriesî; moderate trauma was

defined as ìtrauma that resulted in fatal injuries that could not be determined with certainty to be

instantaneously fatalî; and minimal trauma was defined as ìtrauma in which fatal injuries were present, but

not considered instantaneously fatal.î On the basis of these criteria, investigators determined that 183 of the

202 occupants received severe trauma injuries, 15 of the 202 occupants received moderate trauma injuries,

and 4 of the 202 occupants received minimal trauma injuries. Occupants identified as receiving minimal

trauma injuries were all assigned to seats in zone C of the airplane passenger cabin; two of these four

occupants were recovered from the surface of the ocean during the first hours after the accident. 
Factual Information 87 Aircraft Accident Report

1.16 Tests and Research

1.16.1 Radar Data Information 

The Safety Board reviewed radar data obtained from the FAA (long-range and

airport surveillance radar [ASR]), military, and private sites for evidence of a missile or a

midair collision. Additionally, the Board used radar information, where possible, to locate

and track pieces of the airplane as they fell to the water, to help define the path of the

airplane wreckage. The Board used long-range radar data recorded at FAA radar sites at

Trevose, Pennsylvania; Riverhead, New York; and North Truro, Massachusetts. The

long-range radar systems at each of these sites are capable of recording data within about a

200-nm radius of the site and completing a 360° sweep about every 12 seconds.173 In

addition, the Board used ASR data from five FAA radar sites, which are located at the

following airports: JFK; Newark International Airport, Newark, New Jersey; Long Island

MacArthur Airport, Islip, New York; Westchester County Airport, White Plains, New

York; and Stewart International Airport, Newburgh, New York. Each ASR site is capable

of recording data within about 60-nm radius of the site and completing a 360° sweep about

every 4.7 seconds. The Board also reviewed radar data recorded at the Sikorsky Aircraft

plant in Stratford, Connecticut. These radar data were fed into the Sikorsky plant from the

FAAís Riverhead, New York, radar site and from Sikorskyís own radar site in Shelton,

Connecticut. The Sikorsky radar site is capable of recording data within about a 60-nm

radius of the site and completing a 360° sweep about every 4 seconds. 

At the Safety Boardís request, recorded radar data were also obtained from the

sites of the USAFís 84th Radar Evaluation Squadron (84 RADES). The 84 RADES uses

the same long-range antennas as the FAA (at Trevose, Riverhead, and North Truro) to

gather the raw radar returns. However, the USAF had modified the radar system at North

Truro and used its own processing software, which allowed the 84 RADES to estimate the

altitude of primary targets.174 The North Truro radar system recorded a primary target at a

location that was consistent with the trajectory of the accident airplane just after the last

transponder return from the accident airplane was recorded; this target had an estimated

altitude of 14,400 feet. Because of accuracy limitations,175 this radar data could not be

used to determine whether the accident airplane climbed after the nose separated. (Many

of the subsequent primary returns recorded a value of 102,000 feet, which, according to

the 84 RADES, is a flag denoting that the radar was unable to estimate the height of the

target even though a position was determined.) 

173 The data recorded at these three radar sites were fed to Boston, New York, and Washington, D.C.,

ARTCCs; to U.S. Navy sites; and to a private site operated by Sikorsky Aircraft.

174 At the time of the TWA flight 800 accident, only the North Truro radar system had been modified to

record altitude data for primary returns. 

175 The published accuracy for the recorded altitude values is +/- 3,000 feet; however, the error can

increase significantly when the range of the radar target increases beyond 100 nm, as was the case with the

accident airplane. According to the 84 RADES, the height estimations of higher targets (that is, 20,000 feet

msl and above) were more accurate than the recorded height estimations of targets near the accident

airplaneís altitude. 
Factual Information 88 Aircraft Accident Report

Recorded radar data from ground-based radar antenna sites can be either primary

or secondary in nature. A primary radar target is recorded when a primary radar signal

reflects off of an objectís surface and returns to the site for processing and display.

Primary radar returns contain no identification or altitude information. A secondary radar

target is produced when a radar signal is detected by an airplaneís transponder,176 which

transmits a coded message in response to interrogation by a ground-based transmitter.

Secondary radar returns contain airplane identification and altitude data. All radars

reviewed by the Safety Board during this investigation were transmitting and recording

primary and secondary data, except the Sikorsky radar, which transmitted and recorded

secondary data only. 

Although radar used in ATC is intended to display returns from airplane surfaces,

unassociated primary radar returns (those that are not associated with any airplane track)

are common and can be caused by other surfaces, such as trucks, boats, buildings, weather,

flocks of birds, smoke, temperature inversions, and ground/sea clutter. According to FAA

Order 6300.13, ìRadar Systems Optimization and Flight Inspection Handbook,î chapter

3, paragraph 24, dated March 10, 1992, ì[t]errain, fixed structures, and surface traffic

within line-of-sight range of the ASR antenna system reflect radar energy which can

degrade performance of the system. Such reflections can produceÖfalse target displays.î 

Using the recorded radar data, the Safety Board was able to track TWA flight 800ís

secondary radar returns from the time that the airplane departed JFK (about 2019) until the

time that the airplaneís last secondary radar return was recorded (2031:12). The radar data

review also revealed multiple sets of primary and secondary radar returns from other

airplanes/objects in the area as the accident airplane climbed after takeoff and at the time

of the accident. 

Examination of the radar data showed the following vehicle and/or object tracks

within 10 nm of TWA flight 800 just before the accident (see figure 25):

ï A U.S. Navy P-3 antisubmarine airplane was less than 3 nm south-southwest

of TWA flight 800 at an altitude of about 20,000 feet msl,177 moving to the

southwest at more than 250 knots ground speed. 

ï USAir (now USAirways) flight 217 was about 3 nm south-southwest of TWA

flight 800, descending through an altitude of about 21,700 feet msl and moving

northward. 

176 A transponder is the airborne receiver/transmitter portion of a radar system that responds to

interrogation signals received from ground-based equipment. 

177 Most of the radar returns for the P-3 were primary radar returns because its transponder was

malfunctioning such that secondary signals were available intermittently. During postaccident interviews,

Boston ARTCC personnel stated that they were actively monitoring the P-3ís primary radar returns, despite

the malfunctioning transponder. According to ATC transcripts and postaccident interviews with the P-3

flight crew, the P-3 was at an assigned altitude of 20,000 feet msl when the accident occurred. Further, one

secondary radar return with the P-3ís assigned transponder beacon code was identified among the primary

returns along the P-3ís flightpath, and the 84 RADES primary altitude data showed that the P-3 was about

20,000 feet msl. 
Factual Information 89 Aircraft Accident Report

ï TWA flight 900 was about 9 nm west of TWA flight 800 at an altitude of about

19,000 feet msl, moving to the east-northeast.

ï An unidentified (primary radar) track was recorded less than 3 nm

south-southeast of TWA flight 800, moving southwest about 30 knots ground

speed, consistent with the speed of a boat. 

ï An unidentified (primary radar) track was recorded about 5 nm west of TWA

flight 800, moving east-southeast about 15 knots ground speed, consistent with

the speed of a boat. 

ï An unidentified (primary radar) track was recorded about 5 nm west-northwest

of TWA flight 800, moving to the south-southwest about 12 knots ground

speed, consistent with the speed of a boat. 

ï An unidentified (primary radar) track was recorded about 6 nm northwest of

TWA flight 800, moving to the southeast about 20 knots ground speed,

consistent with the speed of a boat. 

The radar data also showed several isolated primary returns not associated with

any track. (As previously noted, primary radar returns are often recorded from surfaces

other than airplane surfaces.)

The Safety Boardís examination of all of the available radar data revealed no

sequence of primary or secondary radar returns that intersected TWA flight 800ís position

at any time, nor did it reveal any radar returns consistent with a missile or other projectile

traveling toward the accident airplane. No secondary radar returns were received from

TWA flight 800 after 2031:12; however, after 2031:12, numerous new primary radar

returns appeared near the accident airplaneís last recorded radar position, some of which

were visible for up to 20 minutes after the last secondary radar return was received from

the accident airplane. The primary radar returns that appeared near the accident airplane

after 2031:12 were recorded largely in two areas of dense concentration, located about 1 to

1 1/2 miles east-northeast and 1 1/2 to 2 1/2 miles northeast of the last secondary radar

return, respectively (see figure 26). 

1.16.1.1 Information Regarding Certain Primary Radar Targets Recorded by 

the Islip, New York, Radar Site

The Islip, New York, radar data showed one particular sequence of eight primary

radar returns about the time of the accident (see figure 27) that warranted further review.

On the basis of the times of and distances between these primary radar returns, the target

appeared to be moving away from TWA flight 800ís position at a ground speed of more

than 400 knots. The first of the eight radar returns was recorded at 2030:15, more than

6 nm southeast of the accident airplane. The last of these returns was recorded at 2031:30

(18 seconds after TWA flight 800ís last secondary return), about 15 miles southeast of

TWA flight 800. 
Figure 25. Radar data showing vehicle and/or object tracks within 10 nm of TWA flight 800 just before the accident
Figure 26. Primary radar returns that appeared near the accident airplane after 2031:12. 
Figure 27. Sequence of eight primary radar returns recorded by the Islip, New York, radar site about the time of the 
Factual Information 93 Aircraft Accident Report

Although most of the radar systems from which the Safety Board obtained data

were capable of recording radar targets in the geographic area in which these primary

radar returns were recorded, this sequence of eight primary radar returns was only

observed in the data from the Islip radar site. Further examination of the returns recorded

in that geographic area by the Islip site between about 2030 and 2040 revealed two

additional sequences of primary radar tracks with apparent ground speeds greater than

300 knots. One of these sequences indicated a radar target movement away from TWA

flight 800 to the southeast, whereas the other sequence indicated a target movement away

from TWA flight 800 to the southwest. 

In all three cases, the sequences of primary radar targets appeared with no radar

track leading to them and disappeared after four to eight intermittent returns. The returns

in each of these sequences were irregularly spaced and had varying signal strengths. The

azimuth of each of the sequences was 150° to 160° relative to the radar, and there were no

primary or secondary tracks leading to or away from them. None of the three sequences

intersected TWA flight 800ís position at any time. 

Investigators learned from air traffic controllers and radar technicians that ground

or building reflections of primary radar returns from airplanes flying in one geographic

area can cause ìfalse primaryî targets to be recorded as though the airplanes were flying in

another geographic area. FAA documentation and review of the area around the Islip radar

site showed numerous buildings and structures that could create such reflections.

Investigators identified commercial airplanes traveling through other areas within the

coverage of the Islip radar at the same time that the three sequences of primary radar

returns were recorded; these airplanes had similar ground speeds to those indicated by and

flightpaths that were not inconsistent with the three sequences of primary radar returns. 

During its investigation of this accident, the Safety Board became aware of several

allegations of potentially suspicious radar targets, including the following: (1) that the

Islip primary radar data showed evidence of groups of military surface targets converging

in a suspicious manner in an area around the accident site and (2) that a southwesterly

bound 30-knot primary track just south-southeast of the accident site was involved in foul

play, as evidenced by its failure to divert from its course and assist with the search and

rescue operations. 

With regard to the first of these radar targets, the Safety Board examined the U.S.

military records for the time of the accident, which showed that there were no military

surface vessels within 15 nm of the accident site at the time of the accident. In addition,

the U.S. military ìWarning and Restricted Areas Information Logî dated July 17 and 18,

1996, indicated that all warning areas in the area of the accident were available on July 17,

1996, for concurrent use by nonmilitary aircraft (not scheduled for exclusive military use)

except for the following: 

ï Air operations were scheduled in warning area W-387A/B (closest point about

160 nm south of the accident site) between 2030 and 2230. 

ï Air operations were scheduled in warning area W-107 (closest point about

60 nm south of the accident site) between 1330 and 1600. 
Factual Information 94 Aircraft Accident Report

With regard to the second potentially suspicious radar target, the Safety Board

reviewed the 30-knot primary target recorded by the Islip radar site on the night of the

accident to determine why it did not divert from its course and proceed to the area where

the TWA flight 800 wreckage had fallen. The Boardís examination of radar data showed

that at the time of the last TWA flight 800 secondary return (2031:12), the 30-knot target

was about 3 nm south-southeast of TWA flight 800 and headed south-southwest. On the

basis of the targetís position relative to TWA flight 800 and the likely forward-looking

perspective of the targetís occupant(s), the occupants would not have been in a position to

observe the accident airplaneís breakup and/or subsequent explosions/fireball(s).

Additionally, it is unlikely that the occupants would have been able to hear the TWA flight

800 event over the sound of its engines and the ambient noise associated with the hullís

displacement of water. Further, if the occupants were in an enclosed bridge or cabin, the

chance of seeing or hearing the TWA flight 800 event would be even further diminished.

Further, the Board reviewed the Islip radar data for other similar summer days and nights

in 1999 (3 years after the accident) and found numerous slow offshore primary tracks

consistent with those found at the time of the accident. The Boardís review indicated that

such tracks were consistent with normal commercial fishing, recreational, and/or cargo

vessel traffic. 

1.16.2 Trajectory and Main Wreckage Flightpath Studies

1.16.2.1 Trajectory Study Description and Results 

During its investigation of the TWA flight 800 accident, the Safety Board

conducted a trajectory study in which it used the known winds aloft and the weights and

estimated aerodynamic characteristics of selected pieces of wreckage178 to help predict the

likely timing of those piecesí separation from the airplane and their motion during the

subsequent descent to the water. The results of the trajectory study were used in

combination with other information, such as radar data, wreckage recovery locations, and

the metallurgical expertsí examination of the wreckage reconstruction, to help determine

the sequence179 of the airplaneís structural breakup. 

A ballistic objectís180 trajectory is determined by its ballistic coefficient, which is

the weight of an object divided by the product of its drag coefficient multiplied by its

area.181 Thus, a foam ball (which has a very low ballistic coefficient) would fall slowly

when released from an initial point in space, moving almost exclusively with the wind to

its ground location. In contrast, a bowling ball (which has a high ballistic coefficient)

would fall rapidly when released from an initial point, with very little displacement

resulting from the wind. 

178 The trajectory study concentrated on items recovered from the red zone, the westernmost search area

along the wreckage path. 

179 For further information about the Safety Boardís sequencing study, see section 1.16.3.2.

180 For the purposes of this report, a ballistic object is one that has no stable lift. 

181 For specific information regarding individual objectsí ballistic coefficients, see the Trajectory Study

Group Chairmanís report, ìTrajectory Study, Supporting Material,î dated October 14, 1997. 
Factual Information 95 Aircraft Accident Report

The Safety Board used time-step simulation182 to calculate the paths of selected

pieces of wreckage from the airplaneís position at the time of the initial event (based on

FDR data) to that objectís recovery location.183 The Board used its BREAKUP computer

program184 to determine the trajectories of pieces of wreckage that had separated from the

airplane.185 The program begins with the pieces of wreckage at an altitude and airspeed

consistent with the airplaneís last recorded FDR data, then predicts the movement of those

pieces through the air, and generates coordinates for the impact location of these pieces of

wreckage.186

In general, the trajectory study indicated that pieces of wreckage recovered from

the red zone departed the airplane during the first few seconds after the initial event,

followed shortly thereafter by the separation of the forward fuselage (recovered from the

yellow zone). As outlined in the following section, the remainder of the airplane flew for a

time in crippled flight and did not exhibit ballistic behavior until about 40 seconds after

the initial event (when the WCS failed); pieces of wreckage from this portion of the

airplane were recovered from the green zone. In general, the results obtained in the

trajectory study were consistent with other evidence, including the Metallurgy

Structures/Sequencing Groupís (Sequencing Group) findings regarding the airplaneís

breakup.187 However, the trajectory studyís results for two wreckage pieces (RF35 and

CW504)188 suggested that these pieces separated before the airplaneís last secondary radar

return, which was not consistent with other (CVR, FDR, radar data, and sequencing study)

evidence. The Safety Boardís trajectory study indicated that these pieces of wreckage did

not behave ballistically.189

1.16.2.2 Main Wreckage Flightpath Study Description and Results

The Safety Board conducted a series of computer simulations to examine the

flightpath of the main portion of the fuselage (including the wings, most of the WCS, aft

182 Time-step simulation calculates a new position, velocity, flightpath angle, etc. for each piece of

wreckage using accelerations calculated for the previous iteration. 

183 According to Oceaneering personnel, the documented wreckage recovery locations were estimated to

be accurate to within 100 meters. However, the accuracy of documented recovery locations varied

depending on recovery platform and global positioning system (GPS) equipment type and wreckage type.

(Some small pieces were collected in baskets and then lifted to the ship; only the basket location was

documented.) 

184 The BREAKUP computer program is a FORTRAN program developed at the Safety Board that,

given an initial position and velocity vector, calculates the time history of the trajectory of a piece of

wreckage (defined by its drag and weight). 

185 The winds used in the trajectory study were interpolated from upper air data measured by a

radiosonde balloon launched from Upton, New York, on July 17, 1996, at 1934 (see table 3 in section 1.7). 

186 For more details regarding this process, see the Trajectory Study Group Chairmanís report,

ìTrajectory Study,î dated October 14, 1997.

187 For more information, see 1.16.3.2.

188 For additional information regarding CW504 and RF35, see the Trajectory Study Group Chairmanís

report, ìTrajectory Study,î dated October 14, 1997. Also, see the Data Management Group Chairmanís

report, ìData Management Report,î dated November 17, 1997. 

189 See the trajectory study for possible explanations for this nonballistic behavior.
Factual Information 96 Aircraft Accident Report

fuselage, and empennage) after the forward fuselage separated. Investigators configured

the Boardís workstation-based flight simulation computer program to enable simulations

that would reflect the thrust, mass, and longitudinal aerodynamic characteristics of an

intact 747-100, based on Boeingís 747-100 engineering data. Then, investigators modified

the engineering code so that the simulation would reflect the aerodynamic and mass

property changes that Boeing estimated would have occurred when the forward fuselage

separated.190

The separation of the forward fuselage would result in significant changes to the

weight and balance and aerodynamic characteristics of the airplane, substantially

modifying its flight characteristics. Boeingís engineers estimated that the airplaneís c.g.

would move aft from 21.1 to 57.8 percent MAC; the airplaneís total weight would

decrease by almost 80,000 pounds; and the airplaneís moments of inertia would decrease

significantly when the forward fuselage separated. Further, the aerodynamic loads that

would normally result from air impacting and flowing over the smooth forward fuselage

would be replaced by the loads created by air impacting and flowing over the blunt open

fuselage, which would result in increased drag and altered airflow across the inboard

sections of the wings. When the predicted weight and balance changes were considered

with the predicted aerodynamic effects, an initial net nose-up pitching movement was

indicated. 

After the Safety Boardís flight simulation computer program had been modified to

account for the changes in the 747-100ís mass and aerodynamic characteristics,

investigators conducted simulations to identify the airplaneís motion during the seconds

before and after the separation of the forward fuselage. The simulation ìairplaneî was

positioned at the altitude, airspeed, and pitch angles last recorded by the FDR and then

flown through the loss of forward fuselage without further flight control input. The

modified mass and aerodynamic characteristics were applied to the remaining fuselage

until it began to break up,191 after which the remaining pieces of wreckage exhibited

ballistic behavior. Because investigators could not positively determine engine power

settings after the FDR stopped recording, the Board performed simulations using each of

the following engine power settings: (1) the accident airplaneís last recorded power setting

(climb power), (2) engine power off, and (3) full engine power. The results of these

simulations indicated that variations in engine power would not have significantly affected

the path of the crippled airplane.

Rather than develop a simulation for all combinations of aerodynamic

characteristics, the Safety Board bracketed the estimated range of effects by conducting a

ìfastî simulation (using maximum nose-down pitching moment and minimum drag and

lift coefficients), a ìslowî simulation (using minimum nose-down pitching moment and

maximum drag and lift coefficients), and an ìintermediateî simulation (using midrange

190 Because it is not possible (with existing technology) to precisely quantify these changes, Boeingís

engineers estimated a range of values for the airplaneís aerodynamic characteristics after the forward

fuselage separated.

191 Physical evidence described in the sequencing report (see section 1.16.3.2) indicates that the wings

failed and separated outboard of the outboard engines followed by failure of the WCS adjacent to the left

wing.
Factual Information 97 Aircraft Accident Report

data). Further, the Board varied the timing of the forward fuselage separation, conducting

simulations with nose departure times at the beginning, middle, and end of the 4-second

interval indicated for possible forward fuselage separation by the trajectory study. Table 4

shows scenarios involved in the simulations.

The Safety Boardís simulations indicated that the longitudinal response of the

airplane to the loss of the forward fuselage (with no flight control input) was to climb to

above 16,000 feet msl. However, the airplaneís lateral response to the loss of the forward

fuselage (with no flight control input) was less clear. Evaluation of primary radar return

data (from radar sites at White Plains, Islip, and JFK) indicated that after the initial event,

the main portion of the airplane may have turned north after the forward fuselage

separated and then turned south toward the main wreckage recovery site in the green zone.

However, there was considerable scatter in the radar data after the in-flight breakup, and

the recorded radar tracks differed, such that data from the Islip and White Plains radar

sites indicated a more pronounced turn to the north followed by a turn back south, while

data from the JFK radar site indicated a much straighter path to the wreckage site.

Simulations were run using bank angle inputs to match the radar data. The slow

and intermediate simulations that used a nose departure time at the beginning of this

interval did not match the radar data, nor did any of the simulations that used a nose

departure time at the end of this interval. The fast simulation that used a nose departure

time at the beginning of the interval could provide close matches with data from all three

radar sites with different bank angle inputs. Simulations with a nose departure time in the

middle of the interval did not match the Islip and White Plains radar points; however,

intermediate and fast simulations with a nose departure time in the middle of the interval

matched the JFK radar data. (See figures 28a through 28e.) The Safety Boardís

simulations indicated that the maximum altitudes that might have been reached by the

TWA flight 800 aft fuselage after separation of the nose section ranged between 15,537

and 16,678 feet msl. 

In addition to the uncertainties caused by radar track variation, radar data scatter,

and imprecise aerodynamic characteristics previously discussed, there are uncertainties in

the timing and physics of the final breakup that affect the final seconds of the simulations.

The timing of the wing tip and WCS failure were based largely on witness statements;192

therefore, they are not precise to the second. Further, the change in aerodynamics caused

by the wing tip failure at a high AOA is uncertain, as is the effective ballistic coefficient of

the main body after the WCS failure. 

Table 4. A time line of simulation scenarios.

Clock Time(s) Event

2031:12 Initial event

2031:15.2, 2031:17.2, 2031:19.2 Forward fuselage separation

2031:46 Wing tip failure 

2031:50 WCS failure adjacent to left wing

192 For more information, see section 1.18.4 and appendix H.
Factual Information 98 Aircraft Accident Report

Figure 28a. Roll angle for nose off at 2031:15.2 cases.

Figure 28b. Map view of nose off at 2031:15.2 cases.

20:31:00 20:31:10 20:31:20 20:31:30 20:31:40 20:31:50 20:32:00 20:32:10

-180

-160

-140

-120

-100

-80

-60

-40

-20

0

20

40

60

80

100

120

140

160

180

fast simulation nose departure at 20:31:15.2

Simulation for ISP radar

Simulation for JFK radar

Simulation for HPN radar

TWA 800

Main Wreckage Simulation

Time (hr:min:sec)

Roll Angle (deg)
Factual Information 99 Aircraft Accident Report

Figure 28c. Altitude for nose off at 2031:15.2 cases.

Figure 28d. East position for nose off at 2031:15.2 cases. 
Factual Information 100 Aircraft Accident Report

Figure 28e. North position for nose off at 2031:15.2 cases. 

1.16.3 Sequencing Study Information 

In December 1996, the Safety Board formed the Sequencing Group, which

comprised Safety Board investigators and representatives from the parties to the

investigation, to evaluate the sequence of the airplaneís structural breakup and correlate

and compare proposed accident scenarios with the structural observations. To this end, the

Sequencing Group examined individual pieces of the recovered structure,

two-dimensional reconstructions or layouts of sections of the airplane, and various-sized

three-dimensional reconstructions of portions of the airplane.193 The Sequencing Group

also evaluated the relationship of the location of pieces of wreckage at the time of

recovery194 to differences in fire effects (soot deposits on surfaces and fractures, changes

in electrical conductivity) noted on pieces that are normally mated or adjacent to each

other. The following subsections discuss the two- and three-dimensional reconstructions

of the recovered airplane wreckage, the Sequencing Groupís examinations of the

wreckage, and the accident airplaneís breakup sequence identified by the Sequencing

Group. 

193 For additional information regarding the reconstructions, see the Reconstruction Group Chairmanís

Factual Report, dated October 7, 1997; the Medical Forensic Group Chairmanís Factual Report, dated

October 17, 1997; and the Fire and Explosion Group Chairmanís Factual Report, dated August 22, 1997. 

194 For descriptions of most of the pieces of wreckage found in the red, yellow, or green zones, see

sections 1.12.1, 1.12.2, and 1.12.3.
Factual Information 101 Aircraft Accident Report

1.16.3.1 Two- and Three-Dimensional Reconstructions

Several two-dimensional grids were marked on the hangar floor in Calverton,

corresponding to Boeingís drawings of the 747ís various major structures and components

(fuselage, right and left wing structures, cabin interior [including seats, galleys, and

lavatories], empennage, landing gear, and WCS). After the individual pieces of wreckage

were identified and placed in the appropriate location on the corresponding grid, the

structure was examined by the Sequencing Group.

Between August 1996 and January 1997, the WCS/CWT was reconstructed in the

following three separate small three-dimensional mockups: (1) the rear spar, SWB1, mid

spar, SWB2, SWB3, centerline rib, and the WCS upper panel; (2) the front spar, forward

lower panel pieces, keel beam, and adjacent fuselage pieces; and (3) the remaining WCS

lower panel. Investigators also assembled separate three-dimensional reconstructions of

the forward and aft cargo compartments, air conditioning packs, underwing fairing, nose

landing gear wheel well, fragments of fuel tank assemblies and components, and the

cockpit instrument panels. These three-dimensional reconstructions were assembled

informally, with investigators grouping various combinations of parts, in an attempt to

understand spatial relationships between damaged components. Additional documentation

of individual components and pieces of wreckage was conducted while the pieces were in

these reconstructions. 

In the spring of 1997, investigators assembled a 93-foot-long, three-dimensional

reconstruction of the center portion of the airplaneís fuselage (from about STA 520 to

about STA 1632, including the CWT, the main landing gear bay, the surrounding fuselage,

and the furthest inboard pieces of the wings) in the hangar in Calverton.195 Pieces from

smaller three-dimensional reconstructions were incorporated into the large threedimensional reconstruction. Additionally, the Airplane Interior Documentation Group

inventoried, examined, and identified all recovered airplane interior components and

(where possible) placed those interior parts in the interior of the 93-foot long,

three-dimensional reconstruction. This larger three-dimensional reconstruction of the

center portion of the fuselage provided the Sequencing Group with useful information

about damage patterns and failure sequences. (Figure 29 is a photograph of the right side

of the large three-dimensional reconstruction, with the support scaffolding visible.)

Examination of the 93-foot long, three-dimensional reconstruction revealed that the

largest area of unrecovered or unidentified fuselage structure in the vicinity of the CWT

was about 1 square foot. However, the underlying structural pieces (stringers and/or

frames) were recovered in each area of missing fuselage and revealed no evidence of

through penetration. 

195 In December 1996, the Safety Board contracted with Wiss, Janney, Elstner, and Associates, Inc., to

design, fabricate, and assemble the required framework, devise the means by which the various pieces of

airplane wreckage could be attached to the framework, and connect the pieces of airplane wreckage to the

framework. Some truss fabrication and erection tasks were subcontracted to Syracuse Rigging, and steel

fabrication was performed by Delhi Steel Corporation. The reconstruction of this portion of the airplane

occurred between February and late May 1997. 
Figure 29. A photograph of the right side of the large three-dimensional reconstruction, with the support scaffolding visible.
Factual Information 103 Aircraft Accident Report

1.16.3.2 Airplane Breakup Sequencing Study Information 

As previously discussed, pieces of the forward portion of the WCS (including

some pieces from the CWT) and the fuselage directly forward of the wing front spar were

recovered from the red zone and (consistent with the Safety Boardís trajectory study)

were, therefore, among the first pieces to separate from the airplane. Pieces of wreckage

from forward of about STA 840 were found primarily in the yellow zone. Further, most of

the remainder of the airplane (including the wings, major portions of the WCS, the aft

fuselage, and the empennage) was found in a relatively small area in the green zone,

indicating that this portion of the airplane remained intact for some time after the

separation of the forward fuselage section.

The Sequencing Group also conducted detailed visual examinations, occasionally

with magnifications up to 30 X, of the separated structure, with primary emphasis on the

WCS and fuselage pieces recovered from the red zone. Fracture directions, deformations

associated with pieces adjacent to the fractures, and witness marks were evaluated on

many of these pieces of wreckage. The group determined fracture directions based on

visible fracture features that indicated local fracture propagation direction (such as

chevron marks and river patterns [a series of small steps in a fracture that indicate the local

direction of crack propagation]), the presence of branching cracks, and rivet-to-rivet

fracture features.196 Deformation associated with a fracture indicated how the pieces on

each side of the fracture were moving relative to each other as the fracture occurred.197

Witness marks demonstrated the direction of motion of the structure as it separated and

deformed. All of these evaluations provided the Sequencing Group with information about

initial areas of separation. The Sequencing Group also used stress analysis198 to reinforce

that the proposed scenarios for the breakup sequence were consistent with structural

properties and expected failure modes.199

196 Rivet-to-rivet fracture direction was determined by identifying a smaller crack that usually initiated

from a rivet hole in front of the main (larger) crack and then turned to meet the main crack. With fractures

thus divided into longer and shorter legs between rivet holes, the longer leg would indicate the primary

direction of fracture propagation.

197 Fuselage skin fractures that did not contain out-of-plane deformation were classified by the

Sequencing Group as ìhaving occurred earlier in the sequence because these fractures must have occurred

while the associated fuselage pieces were still aligned with each other nearly in the correct position.î

Fuselage skin fractures that did contain out-of-plane deformation (either bending or tearing) were classified

by the group as ìhaving occurred later in the sequence because the associated pieces of fuselage must have

been out of alignment with each other before the fracture occurred.î Fuselage skin fractures along a row of

rivet holes that were consistent with in-plane loading (hoop or longitudinal tension with the possible

presence of in-plane shear) were classified by the group as ìhaving occurred earlier in the breakup sequence

than fractures with out-of-plane bending or shear deformation.î 

198 Stress analysis is the science of using calculations to estimate stress. Inputs into the calculations

include information on, for example, material properties, component dimensions, structural conditions, and

load amounts and directions. 

199 In addition, Boeing calculated failure strengths for the various CWT components (spars and beams)

and the pressures at which the CWT components would fail and provided the results to the Safety Board. For

more information about the these calculated failure strengths and pressures, see section 1.16.4.9. 
Factual Information 104 Aircraft Accident Report

The Sequencing Group examined portions of the structure in great detail and then

developed localized sequence segments based on the observable features in each portion.

The group agreed that any proposed sequence for an individual portion of the structure

had to comport with the general breakup sequence indicated by the overall recovered

structure, including the early separation of forward portions of the WCS and a ring of

fuselage material from forward of the WCS. Individual sequence segments were

combined, until a cohesive and comprehensive breakup sequence, consistent with the

overall body of evidence, was generated.

The detailed examinations of the fuselage red zone pieces are an example of how

these examinations were successful in determining an important segment of the breakup

sequence. Figure 30 shows a schematic drawing of the lower portion of the fuselage skin

in the area forward of the front spar. In this figure, the earlier fuselage skin fractures have

been highlighted, and arrows have been added to indicate the direction of cracking in these

earlier fractures. Of all of the fuselage fractures examined, only a limited number (all of

which were on pieces recovered from the red zone) could be classified as earlier fractures

(that is, consistent with in-plane loading). Further detailed examination of the rivet-torivet fracture pattern in these earlier fuselage fractures generated information on crack

propagation directions and indicated that the earliest fracturing was of the lower fuselage,

beginning at the front spar of the WCS. (The methodology used to determine the sequence

of fracture in the fuselage red zone pieces was typical of the approach used for the entire

airplane structure.) 

On the bases of its detailed examination of the recovered airplane structure (which,

as previously indicated, considered fracture features and other physical evidence, recovery

locations, etc.),200 the Sequencing Group determined that the initial failure of the airplane

structure was the spanwise fracture of SWB3 at its upper chord. The sequencing report

identified an overpressure event within the CWT as the cause of the fracture of SWB3.

According to Boeingís calculations, SWB3 is the weakest of the boundary members of the

CWT and would be the first member expected to fail as pressure within the tank increased.

As discussed in the next several paragraphs, evidence of the overpressure event within the

CWT was apparent in the upward lifting of the center of the CWT upper skin as SWB3

failed at its upper chord, in the subsequent forward rotation of SWB3, and in the forward

bowing of the front spar as it fractured. 

As the upper portion of SWB3 rotated forward, hinging on the still-attached lower

chord, the upper chord impacted the aft side of the upper skin panel stiffener immediately

forward of SWB3. (Figure 31 shows the approximate shape of this stiffener and its

location before being struck by SWB3 as the beam rotated forward.) A geometrical layout

of SWB3 and the upper and lower skin panels indicates that the forward rotation of SWB3

about its lower chord would result in an impact on the vertical flange of the stringer about

1.8 inches above the lower edge of the stringer (assuming no relative motion between the

upper skin panel and SWB3). However, in the center of this stiffener, a witness mark was

200 See Metallurgy/Structures Sequencing Group Chairmanís Factual Report Sequencing Study, dated

April 8, 1997. 
Factual Information 105 Aircraft Accident Report

Figure 30. A schematic drawing of the lower portion of the fuselage skin in the area 
Factual Information 106 Aircraft Accident Report

noted at a distance of only 0.9 inch above the lower edge of the stringer. The difference

between these values represents an upward vertical displacement of 0.9 inch between the

upper and lower skin panels along the airplaneís centerline when SWB3 impacted the

stringer. The distance of the witness mark from the bottom of the stringerís vertical flange

gradually increased at positions further outboard, indicating lesser amounts of upward

vertical displacement outboard of the centerline. According to the sequencing study, the

location of the marks left on the stiffener by SWB3 indicate that the upper skin of the

CWT was likely bowed upward, apparently also as a result of an overpressure event

within the CWT. 

Continued forward and downward rotation of the top of SWB3 caused the upper

end of SWB3 to impact the stiffeners on the aft surface of the front spar about 12 inches

below the upper skin panel. This impact left a distinct set of marks on the front spar and

initiated multiple failures near the top of that spar. (Figure 31 also shows the approximate

orientation between the front spar and SWB3 as the aft side of the front spar was impacted

by the upper end of SWB3.) Geometric evaluation and physical evidence indicated that

the front spar was intact and in place when it was contacted by the upper end of SWB3.

(The location of these SWB3 impact marks on the front spar was a major factor in the

Sequencing Groupís determination that fracture of the upper portion of SWB3 was the

initial failure in the airplane structure.) 

According to the sequencing study, impact of SWB3 with the aft side of the front

spar severely damaged the front spar stiffeners and initiated fractures in the front sparís

upper chord. Overpressure escaping from the CWT caused the front spar to bulge forward

as it failed along the upper chord; this bulging resulted in two forward-protruding lobes on

the front spar, one on each side of the spar centerline. The sequencing report attributed the

formation of the two lobes to the inertial resistance provided by the two potable water

bottles attached to the center of the forward side of the front spar. These bottles were full

when the airplane left JFK, and their combined weight was more than 3,000 pounds. As

the front spar bulged forward, fractures developed near the center of each of the two lobes.

In addition, the front spar was further damaged when portions of SWB3 penetrated the aft

side of the front spar as SWB3 continued to rotate forward and down. 

The Sequencing Group determined that the upper chord of the front spar then

completely separated from the WCS upper skin panel, and the overpressure within the

CWT forced the WCS lower skin panel and the forward end of the keel beam downward.

Stress analysis performed as a part of the sequencing study showed that the continuing

downward keel beam motion increased the stresses in the ring chord and fuselage skin

adjacent to the front spar. The front spar stiffener at LBL 18 was not bent as far forward as

the others, indicating that there was limited forward rotation of the front spar in this area

before the ring chord separated at the bottom of the stiffener. Fractures on the right side of

the front spar continued into the lower pressure bulkhead,201 then continued through the

ring chord and into the fuselage skin just in front of the front spar, at stringer 40 right (see

figure 30). 

201 The lower pressure bulkhead extends between the lower chord of the front spar and the ring chord. 
Factual Information 107 Aircraft Accident Report

Front Spar

SWB 3

(Rotated Forward)

WCS Lower Skin

Top of

Intercostal

STA

990,42

WL

17.78

Approximate

location of witness

marks on the aft

side of the

upper skin panel

stiffener located

just forward of

the upper end

of SWB3

Location of witness

marks on the

aft side of front

spar stiffeners

Initial Position

of SWB3

Figure 31. Diagram showing the approximate shape of the upper skin panel stiffener 

and its location before being struck by SWB3 as the beam rotated forward. 
Factual Information 108 Aircraft Accident Report

On the basis of cracking directions of the fuselage skin, the Sequencing Group

determined that the initial fuselage skin cracking (at stringer 40 right just forward of the

front spar, about STA 1000) propagated forward and toward the bottom center of the

airplane, then continued circumferentially in both directions and back toward the front

spar on the lower left side. This fuselage cracking then quickly progressed202 around three

sides of a large piece of fuselage belly structure (subsequently identified primarily as piece

LF6A). See figure 30 for a diagram of this cracking in the fuselage; piece LF6A is located

in the center of this figure. 

According to the sequencing study, the combined load of normal cabin

pressurization and vented overpressure from the CWT/WCS generated a downward force

on this large section of the fuselage belly structure and resulted in additional downward

loading on the keel beam. Stress analysis showed that these loads were sufficient to cause

separation of the forward portion of the keel beam from the WCS lower skin panel and

then to cause the keel beam to fracture about 22 inches aft of the mid spar. The report

indicated that when the large piece of fuselage belly structure (primarily piece LF6A)

separated, a large opening in the bottom of the fuselage resulted. Pieces from the WCS

(primarily pieces from SWB3, the front spar, and the manufacturing access door from

SWB2) exited the airplane through this large opening. 

As previously discussed, SWB2 contains a removable maintenance access door

(located just to the left of the SWB2 centerline) and a ìpermanentlyî sealed

manufacturing access door (located just to the right of the SWB2 centerline). The

Sequencing Group noted that the manufacturing access door (with an attached piece of the

web of SWB2) was recovered from the red zone and identified the following three

possible reasons for the separation of this door early in the accident sequence: (1) internal

pressure generated as a part of the initial overpressure event created a load that resulted in

the separation of the door; (2) the downward movement of the forward portion of the keel

beam as it separated from the lower surface of the WCS created a downward loading on

SWB2, resulting in separation of the door; or (3) a combination of (1) and (2). 

The sequencing study indicated that, as the belly structure separated, adjacent

pieces of the remaining fuselage skin and structure continued to crack and tear. Nearly

symmetric pieces of fuselage skin above the right side and left side of the large hole in the

belly separated from the rest of the airplane in an outward, upward, and aft direction. A

curl of metal was created on both of these two pieces at the final point of separation (the

upper aft corner). After these pieces of fuselage skin separated, the hole in the fuselage

extended across the entire bottom of the airplane between the main cabin window belts. 

According to the Sequencing Group, as the airplane continued to depressurize

through the large hole in its underside, the nose of the airplane began to bend down,

202 Although the Sequencing Group did not determine specific timing for the events that it identified,

explosive decompression testing showed that normal cabin pressure differentials would cause this type of

crack in the fuselage skin to propagate in fractions of a second. (The experts involved in Fire and Explosion

Group testing calculated that this fuselage cracking occurred in about 24 milliseconds.) In addition to

normal cabin pressurization loads, the accident airplaneís fuselage structure was subjected to loads from the

venting of the WCS overpressure. 
Factual Information 109 Aircraft Accident Report

creating compression stresses, in the window belts above the hole. The window belts

collapsed from these compression stresses and compression buckling spread upward

toward the crown of the airplane, where evidence of tension failure was found.

Subsequently, pieces of fuselage skin began to separate from right to left across the top of

the airplane; some of these pieces had curls similar to the two symmetric curls below the

window belt. This loss of fuselage structure around the airplaneís circumference forward

of the WCS resulted in the separation of the airplaneís forward fuselage from the

remainder of the airplane, which included most of the WCS, the wings, the aft fuselage,

and the empennage. 

According to the Sequencing Group, fire damage and soot deposit patterns on

recovered wreckage indicated that some areas of fire existed on the main portion of the

airplane as it continued in crippled flight after the loss of the forward fuselage.203 The

study indicated that, after some period of crippled flight (about 34 seconds, based on

information from witness documents), the right and left wings outboard of the outboard

engines failed in upward bending overload. Next, the WCS failed in upward bending

adjacent to the left wing. (The upper skin panel buckled in compression, and the lower

skin panel failed in tension.) The compression buckling in the left side of the WCS upper

skin panel continued aft through the left side of the horizontal pressure deck above the

wing landing gear wheel well and into the body landing gear wheel well. The main

landing gear beam was separated in upward bending below this area of compression

buckling (about LBL 75), consistent with the upward bending failure of the WCS. 

The Sequencing Group determined that the external and internal soot deposit

patterns revealed a distinct difference between structure that separated with the left wing

(as a result of the upward bending failure of the WCS) and the remainder of the airplane.

Lack of soot deposits on the large fuselage piece above the left wing and other fuselage

pieces from the crown of the airplane indicated that these pieces separated with the left

wing. A review of soot deposit patterns, fracture characteristics, and interface of the rear

spar with upper and lower skin panels (see the earlier discussion in this section regarding

evidence from the reconstruction of the remaining WCS upper panel) indicated that the

pickle fork fitting204 on the left side of the rear spar remained attached to the left wingís

lower chord and that the portion of the rear spar located to the right of LBL 21.5 remained

attached to the right wing. Soot deposit patterns on the remaining right wing and aft

fuselage pieces indicated that these pieces of the airplane remained together for some time

after the left wing separated. However, damage characteristics and differences in soot

deposit patterns indicated that the remaining aft fuselage separated from the right wing in

stages.205

The sequencing study stated that upward crushing damage found on the lower

portion of the aft fuselage (aft of about STA 1480) indicated that most of this portion of

203 This information is also consistent with observations recorded in witness documents

(see section 1.18.4) and the Fire and Explosion Group Chairmanís Factual Report, dated August 22, 1997.

204 The rear spar body bulkhead is commonly termed the ìpickle fork fitting.î The pickle fork fitting is a

complex machined forging that provides one of the primary vertical load paths connecting the spars of the

outboard wings and the WCS to the fuselage structure. 
Factual Information 110 Aircraft Accident Report

the fuselage struck the water as a large single piece. On the basis of the location of the

crushing damage, the Sequencing Group determined that the aft portion of the fuselage

was rolled slightly to the right and relatively flat when it impacted the water. Similar

crushing damage was found on the lower portion of the fuselage on the nose section

(forward of about STA 840). The location of the crushing damage indicated that this

portion of the airplane was rolled about 45° to the right when it impacted the water. As

previously discussed, severe crushing damage was not found on any of the pieces found in

the red zone. 

According to the sequencing study, some aft fuselage pieces and wheel well

structure (including the aft portion of the keel beam, the right main landing gear beam, and

the right portion of the STA 1350 bulkhead) remained with the right wing for a period of

time,206 accumulating additional soot from a fire source on the right wing, after the

separation of the fuselage aft of STA 1480. Several right-side fuselage pieces contained

heavy soot deposits and fire damage, indicating that they remained attached to the right

wing during an intense fire. A right-side fuselage piece that was recovered attached to the

right wing was severely burned. 

According to the sequencing study, severe fire damage on many portions of the

WCS that were not recovered from the red zone (including the far right side of SWB3, the

right two-thirds of SWB2, the right portion of the WCS upper skin panel, and the rear spar

near BL 0) indicated that these areas remained attached to the right wing during the most

intense fire. Although very heavy soot deposit and severe fire damage were noted on the

exterior surfaces of the right rear portion of the WCS, the interior surfaces in this area

(including the aft portion of the right side-of-body rib) did not contain fire damage or soot

205 Before, during, and after the Sequencing Groupís study of the airplaneís breakup sequence, Boeing

conducted separate analyses of several portions of the breakup sequence using a computer model of the

747-100 fuselage from STA 540 to 1480, which included most of the WCS box. The computer model was

adapted to simulate the failure of SWB3 and the front spar as described in the previous section. For the

purposes of these analyses, a sustained overpressure of 25 pounds per square inch (psi) was assumed in the

WCS. According to Boeing, it used 25 psi because it was somewhat higher than the pressure at which

Boeing personnel would expect the SWBs to fail (see section 1.16.4.9). The fuselage was pressurized to

4 psi cabin pressure differential. 

Boeing also performed wing bending moment analyses to determine the effects of the separation of the

forward section of the fuselage. These analyses assumed that the aft fuselage and wings would pitch up as a

result of the sudden aft movement of the c.g. Boeingís analyses showed that at a high AOA and a 5.5 to 6 G

load factor, the wing tips would fail, and the WCS would continue to carry the predicted loads. When the

aerodynamic loading assumptions in the analyses were modified to account for the wing tip separation, the

simulation showed that an additional 1 G increase in load factor would result in buckling (and ultimate

failure) of the upper skin panel of the WCS, consistent with the upward bending failure of the WCS

described by the Sequencing Group. According to Boeingís wing bending analyses (wing tip on and wing tip

off), SWB2 must have been sufficiently intact (after separation of the forward fuselage) to provide

substantial support to the WCS upper skin panel. 

According to a January 20, 2000, letter from Boeing, its computer modeling and wing bending

moment analyses did not reveal any inconsistencies with the basic findings described in the Safety Boardís

sequencing report. 

206 The Sequencing Group was unable determine how long each of these pieces remained with the right

wing as the pieces of wreckage fell to the water. 
Factual Information 111 Aircraft Accident Report

deposits. The Sequencing Group determined that this was evidence that the right rear

portion of the WCS remained largely intact until water impact.

The sequencing study established that the most intense fire effects were limited to

the right wing, portions of the WCS, and a limited amount of fuselage structure above the

right wing and that the most likely source of fuel for this fire was the right wing inboard

fuel tank. Many of the fractures from the portion of the airplane with the most intense fire

effects did not contain soot deposits, indicating that significant breakup damage occurred

as the intensely fire-damaged structure impacted the water. 

In summary, according to the sequencing study report, the breakup sequence was

as follows:

ï An overpressure within the CWT initiated a failure at the top of SWB3.

ï The top of SWB3 rotated forward and down and impacted the aft side of the

front spar.

ï The front spar bulged forward in two lobes, one on each side of the potable

water bottles mounted in the center of the forward side of the front spar.

ï Cracking on the right side of the front spar progressed downward, through the

lower-pressure bulkhead and ring chord, and entered the fuselage. 

ï Fuselage cracking propagated dynamically around a large piece of belly

structure, which separated, creating a hole through which WCS structure was

ejected from the airplane. 

ï The forward portion of the keel beam was peeled off of the lower surface of the

WCS, and the keel beam separated in two pieces aft of the mid spar.

ï Fuselage pieces on each side of the hole in the belly structure began to separate

upward, outward, and aft. 

ï The window belts above the hole in the lower fuselage buckled, and this

compression buckling spread toward the top of the airplane. 

ï Pieces of fuselage separated across the top of the airplane, thereby completely

separating the forward fuselage from the remainder of the airplane. 

ï The main portion of the airplane (including the wings, most of the WCS, and

the aft fuselage) continued in crippled flight for a period of time.

ï The wing tips separated at the outboard engines.

ï The WCS separated adjacent to the left wing.

ï The fuselage aft of STA 1480 separated from the remainder of the WCS and

the right wing.

ï Other structure separated from the right wing and remaining portion of the

WCS. 

The Sequencing Group also examined the reconstructed portion of the airplane in

areas where structure appeared to be missing. For example, an apparent hole (2 to 3 feet

long with about a 5 feet circumference) was visible in the fuselage structure above the
Factual Information 112 Aircraft Accident Report

main portion of the pickle fork fitting at the upper left rear corner of the WCS. However,

the Sequencing Groupís report indicated that examination of the fuselage skin surrounding

the apparent hole revealed that most, if not all, of the fuselage skin that appeared to be

missing was present in folded skin attached to adjacent structure. All of the structural

pieces in the vicinity of this ìholeî were recovered from the green zone, indicating that

none of the pieces from this area separated early in the breakup sequence. The Sequencing

Group determined that the features of the apparent hole were consistent with the

compression buckling failure of the left side of the WCS associated with the loss of the

left wing later in the breakup sequence. The Sequencing Group report did not associate

any of the apparent large holes in the structure with the initiation of the airplane breakup.

In addition, the sequencing study report stated that ìone noteworthy aspect of the

structural breakup characteristic is the exceptional degree of symmetry between right and

left sides. Starting with SWB3 and forward in the sequence there is the suggestion of a

very uniform driving force.î 

1.16.4 Additional Metallurgical, Material, and Structural Testing 

and Information 

1.16.4.1 Metallurgical Examination of Several Small Holes in the Accident 

Airplaneís Structure

The Safety Board examined the recovered wreckage for evidence that the

airplaneís breakup was initiated by penetration of the CWT by a small, high-velocity

fragment (generated, for example, by explosion of a missile warhead near the airplane or

of a high-energy explosive within the airplane). Members of the Structures Group

identified 196 relatively small holes in the accident airplaneís structure that they

considered to be in an appropriate location and of an appropriate size and shape to have

resulted from such a fragment. These holes were then examined in detail by Safety Board

metallurgists and others to determine if they had characteristics of high-velocity

penetrations. 

At the Safety Boardís request, Boeing created test plates with a variety of highand low-velocity penetration characteristics, which were then compared with the holes in

the accident airplane wreckage. To produce these test plates, Boeing personnel fired

fragments of various sizes and materials (to simulate missile warhead fragments) at target

plates of various thicknesses and materials (corresponding to those found in the accident

airplaneís WCS). 

Boeing also measured each fragmentís velocity when it impacted the test plate.

Examination of the holes and impact marks on the test plates revealed that holes resulting

from high- and low-velocity penetrations each had distinctive damage characteristics. The

approximate impact velocity at which the features transitioned from high- to low-velocity

(as defined below) was a function of the thickness of the impacted material and the size

and material of the impacting fragments. For example, it was determined that 1/4-inch to

3/8-inch steel cubes impacting a 0.10-inch-thick aluminum plate would produce

low-velocity characteristics at impact speeds less than or equal to 1,000 feet per second
Factual Information 113 Aircraft Accident Report

(fps) and would produce high-velocity characteristics at impact speeds greater than or

equal to 1,500 fps. Characteristics of high-velocity penetrations were as follows:

ï Splashbackóhigh-velocity penetrations caused material around the perimeter

of the hole on the entry side of the plate to deform in the direction opposite to

the motion of the penetrating fragment. This feature occurred for high-velocity

impacts regardless of whether they resulted in full penetration, partial

penetration, or just impact marks. 

ï Exit deformation or breakoutósimilar to splashback but on the exit side of the

test plate. Passage of the fragment through the test plate generated local

deformation or shear breakout of material around the hole. 

ï Hole wall characteristicsóhigh-velocity penetrations created hole walls that

exhibited melted and resolidified material adjacent to the entry side (including

transferred material from the penetrating fragment), severe shear deformation,

and multiple shear cracks. 

ï Overall deformationóhigh-velocity penetrations did not produce significant

overall membrane deformation in the direction of motion of the penetrating

fragment. 

Characteristics of low-velocity penetrations were as follows:

ï Splashbackólow-velocity penetrations did not cause splashback. 

ï Hole wall characteristicsólow-velocity penetrations created hole walls that

appeared to be generated by fracture of the test plates, resulting in elongated

ductile dimples throughout the entire thickness of the plate. No melting was

found on the low-velocity holes. 

ï Overall deformationólow-velocity penetrations produced more overall

membrane deformation in the direction of motion of the penetrating fragment. 

An initial evaluation showed that many of the holes in the recovered airplane

wreckage were obvious penetrations from the inside of the structure outward, contained

substantial deformation adjacent to them, and generally had some amount of impact

damage associated with them consistent with low-velocity penetration by a structural

member. Other holes were actually tears or penetrations in the structure with minimal

associated missing material. On the basis of these observations, all but 25 of the 196 holes

were determined not to have characteristics of high-velocity penetration. The Safety

Boardís more detailed examinations of the remaining 25 holes revealed that 23 of the

holesóall associated with the WCSódid exhibit characteristics of low-velocity

penetration. These holes exhibited no evidence of splashback, and significant deformation

was observed in the surrounding material. 

Only two holes had any characteristics of high-velocity penetration, and these

holes also exhibited some low-velocity penetration characteristics. Both holes were about

3/16 inch in diameter and were located in the horizontal pressure deck above the main

landing gear bay (aft of the WCS/CWT). The hole in piece LF137 was located in the

pressure deck at STA 1457 and LBL 110. The chipped-out metal on the lower surface of
Factual Information 114 Aircraft Accident Report

and the exit deformation associated with this hole were consistent with penetration by a

fragment moving downward and slightly aft. The hole in piece RF60 was located in the

pressure deck at STA 1452 and LBL 62. The chipped-out metal on the lower surface of

and the exit deformation associated with this hole were consistent with penetration by a

fragment moving downward and slightly inboard. Both holes exhibited the following

features:

ï lack of overall deformation in the sheets around the holes;

ï no splashback;

ï chipped-out metal on the lower surface of the sheets on one side of the holes;

and

ï smooth hole walls, generally perpendicular to the surface, with some exit

deformation on the lower surfaces. 

Because these holes contained features characteristic of both low-velocity holes

(no splashback) and high-velocity holes (lack of overall deformation), it was difficult to

determine the likely impact velocity in that area. However, on the basis of the direction of

the penetrations and their locations relative to each other, it was determined that the

penetrating objects for the two holes did not originate from a common point, nor did they

originate from a point outside the airplane. 

A portion of piece LF137 that contained one of the holes was cut from the piece

and sent to the Safety Boardís Materials Laboratory for further examination.207 A bench

binocular microscopic examination (using magnifications of up to 70 X) of this portion of

piece LF137 revealed a crescent of deformed material on the upper surface of the piece,

adjacent to one side of the hole. The hole was slightly elongated in the direction of the

crescent; the hole diameter at the crescent was 0.19 inch, whereas the perpendicular

diameter was 0.17 inch. Examination of the hole using a scanning electron microscope

(SEM) revealed that much of the hole wall was covered by deposits that were mostly

composed of mud-covered oxides. Portions of the hole with fewer deposits within the

deformed crescent exhibited parallel lines consistent with metal deformation smearing.

EDS examination of various areas on the removed piece generated spectra typical of

7075-T6 aluminum alloy,208 the specified material for the piece. In some places, spectra

showed traces of oxygen, sulfur, and chlorine. No evidence of melted and resolidified

metal was noted on any portion of the hole wall. 

1.16.4.2 Metallurgical Examination of Fatigue Cracking 

During their examination of the recovered pieces of the WCS, investigators

observed evidence of several fatigue cracks.209 All of the fatigue cracks detected during

structural examinations were evaluated by the Sequencing Group as a part of the

207 The hole in piece RF60 was directly adjacent to the steel framework supporting the reconstructed

airplane and was much more difficult to cut out without removing the entire piece. Therefore, because the

holes in pieces RF60 and LF137 were so similar in appearance, only the hole in LF137 was examined in the

laboratory. 

208 7075-T6 is an aluminum alloy with primary alloying elements of zinc, magnesium, and copper. 
Factual Information 115 Aircraft Accident Report

determination of the breakup sequence. Small preexisting fatigue cracks were observed in

the front spar vertical stiffener shear ties at RBL 83.24 (lower), RBL 75.92 (upper and

lower), LBL 75.92 (upper and lower), and LBL 83.24 (lower). These cracks were all

located in the shear tie radius near the base of the leg that attaches to the vertical stiffener

near the aft edge and had a maximum length of 0.20 inch. These shear ties are subject to

inspection per Boeing SB 747-57-2249.210

Small fatigue cracks were also found in the shear tie of the LBL 75.92 and

LBL 33.99 longitudinal floorbeams where they intersected the front spar upper chord at

STA 1000.211 The LBL 75.92 shear tie had a 0.15-inch fatigue crack emanating from the

aft side of the bolt hole and a possible 0.125-inch fatigue crack emanating from the

forward side of the hole. The LBL 33.99 shear tie had a 0.25-inch fatigue crack emanating

from the forward side of the bolt hole. 

Additionally, fatigue cracks were found in the lower chord of the front spar. The

cracks were in the horizontal leg of the chord and initiated from the chordís fillet radius

just outboard of the underwing longeron splice fittings at both RBL and LBL 80. The

fatigue crack at RBL 80 was about 1.2 inches long and about 0.10 inch deep, whereas the

fatigue crack at LBL 80 was about 1.45 inches long and about 0.12 inch deep (about 1/3 of

the chordís thickness). 

This region of the accident airplane had been modified in 1982 (in accordance with

SB 747-53-2064, which addressed adjacent ring chord cracking) by incorporating two

fittings on the WCS lower skin panel and a corresponding double fitting on the fuselage

skin. These fittings are adjacent to the underwing longeron splice fitting and provide an

alternate load path for the longeron forward/aft loads. The sequencing report noted that the

postmodification configuration was very stiff and significantly limited further deflection;

without continued deflection, fatigue crack growth cannot continue. Therefore, the

Sequencing Group determined that the fatigue cracking observed in the recovered pieces

of the accident airplane existed before these fittings were installed. The Sequencing Group

further noted that the propagation of the overstress fracture through the fillet radius during

the overpressure event coincidentally exposed these two areas of localized, preexisting

fatigue cracking in the front spar lower chord.

209 Although the accident airplaneís structure contained minimal preexisting corrosion damage, the

Sequencing Groupís study indicated that this damage would not have affected the breakup of the airplane.

Further, the Sequencing Group determined that none of the small fatigue cracks that were found on some

parts of the airplane (as discussed in this section) had coalesced into a propagating crack that could have led

to the in-flight breakup.

210 According to Boeing personnel, the company has received reports from 747 operators observing

cracks ranging from 0.50 to 1.5 inches long without complete part failure. 

211 According to the sequencing report, the front spar upper chord at STA 1000 is a secondary attachment

for the floor structure and was not involved in carrying primary airframe loads. 
Factual Information 116 Aircraft Accident Report

1.16.4.3 Metallurgical Examination of the Fuselage Joint Between Sections 

41 and 42

The Safety Board was aware of concern about possible assembly problems with

the fuselage joint between sections 41 and 42 (which extend forward and aft, respectively,

from STA 520) on 747s212 and considered the possibility that a failure at this joint might

have occurred, initiating the breakup sequence. This joint area was wholly contained

within the forward fuselage section that the Sequencing Group identified as having

separated from the airplane early in the breakup sequence. 

Overall examination of the forward portion of the airplane showed that sections 41

and 42 exhibited uniform crushing damage that extended across the bottom of the fuselage

from about 3 feet below the left-side main cabin windows (S-39L) to above the right-side

main cabin windows (S-14R). This damage was consistent with this section of the airplane

impacting the water intact in a right-wing-low attitude. The structures examination also

revealed that portions of the joint between sections 41 and 42 were separated during the

water impact, whereas some portions remained intact.213 The portions of the STA 520 joint

that separated were examined for evidence of preexisting fatigue or other preexisting

damage. The Safety Boardís examination of this joint revealed that all fractures were

typical of overstress separations, with no evidence of fatigue, fretting, or significant

corrosion. 

1.16.4.4 Metallurgical Examination of the Forward Cargo Door

The Safety Board also considered the possibility that the forward cargo door (the

forward edge of which is located several feet aft of STA 520 on the lower right side of the

fuselage) separated from the accident airplane in flight and that this separation initiated

the breakup sequence. The Board examined the pieces of the forward cargo door, which

were recovered from the yellow zone. All eight of the latching cams at the bottom of the

door were recovered attached to pieces of the lower end of the door and were in the

latched position. Additionally, the latching cams and pieces of the cargo door remained

attached to the pins along the lower door sill. The hinge at the top of the door was broken

into several pieces, but the hinge pin still held the various pieces of the hinge together.

There was no evidence to suggest that this hinge separated. The forward cargo door

exhibited severe crushing deformation and fragmentation, very similar to damage

observed on the adjacent fuselage structure. 

212 In conversations with Safety Board staff, Boeing personnel indicated that because the shape and size

of the ends of sections 41 and 42 could vary during the manufacturing process, extra effort was sometimes

required to fit these sections exactly together. 

213 The fuselage joint between sections 41 and 42 was not separated across the top of the airplane

(between S-2L and S-5R) or along portions of the right and left sides (between S-18L and S-39L, S-44L and

S-47L, S-19R and S-28R, and S-40R and S-45R). 
Factual Information 117 Aircraft Accident Report

1.16.4.5 Metallurgical Examination of Nose Landing Gear Doors and 

Surrounding Structure 

The left aft nose landing gear door was heavily damaged in a manner consistent

with other parts recovered from the yellow zone and with many other pieces from the

forward portion of the fuselage. However, the right and left forward and the right aft nose

landing gear doors exhibited relatively little damage and were recovered from the red zone

with other airplane pieces that the sequencing report indicated had separated early in the

breakup sequence. Further, the hinges from some of these doors appeared to have

overtravel damage, indicating that the doors may have been forced inward. After careful

examination of the nose landing gear doors and pieces of the surrounding structure, the

Sequencing Group determined that the doors did not overtravel inward a significant

amount and that the apparent overtravel damage to the hinges could have occurred after

the doors opened or separated. Subsequent teardown inspection of the nose landing gear

door actuator revealed that the actuatorís locking mechanism was not damaged and was

within specifications, consistent with the Sequencing Groupís failure scenario that the

nose landing gear doors may have opened during the initial stages of the breakup and then

been torn off by exposure to the air stream.214

1.16.4.6 Examination of Brown Splatter Material 

As previously discussed, parts of the WCS and portions of the cabin interior above

and forward of the WCS were found to have splatterlike deposits of a dark brown rubbery

(spongy) material. The highest concentration of these deposits appeared on the top surface

of the WCS near a long duct that was fastened to the top surface of the CWT. Samples of

this brown splatter material were examined, the features of the deposits were recorded in

detail, and their composition and physical characteristics were determined by chemical

analyses. At the Safety Boardís request, several independent laboratories performed

analyses on this splatter material.215

Chemical analyses by the laboratories were consistent and showed that the dark

brown splatter material was consistent with a polyurethane elastomer. This material was

consistent with the flexible polyurethane material with which the long air conditioning

duct located above the CWT was insulated. However, microscopic examination of the

splatter material showed that some of the deposits were not uniform and had fiberlike

materials embedded in the dark material. According to the splatter report, the other

embedded materials or fibers were consistent with a ìNylonô 6 series. The discolored

fiberÖappeared to be a fibrous material much like Azlon.î216 Examination of a long air

214 Cable and hydraulic lines for the nose landing gear control system are routed through the CWT area.

It is likely that these systems would be disrupted early in the airplaneís break up sequence, resulting in a loss

of hydraulic power. 

215 For additional information about the splatter material, splatter pattern and chemical analyses, see

Factual Report on Splatter Deposits, dated October 2000. 

216 According to the Federal Trade Commission, Azlon is a generic name for a manufactured fiber in

which the fiber-forming substance is composed of any regenerated naturally occurring protein (for instance,

proteins from corn, peanuts, and milk have been used in such manufacturing applications). Azlon is

commonly used in airplane carpeting and seat fabrics. 
Factual Information 118 Aircraft Accident Report

conditioning duct removed from the same area on a 747 airplane of similar age revealed

the presence of similar fibers adhering to the outside surface of the duct. These fibers were

determined to be consistent with the fibers in the carpeting and seat fabrics from that

airplane. 

1.16.4.7 Examination of Red/Brown-Shaded Substance Found on 

Passenger Seats 

During the Safety Boardís examination of recovered wreckage, investigators

observed red/brown-shaded substances on several damaged passenger seats. These

substances appeared as a thin coating, which covered the surfaces of interior components

of the seatbacks. Specifically, these substances were found on the forward face (inside) of

the plastic seatback, on the backward face (backside) of the metal seatback frame, and on

the plasticized foam strips used to attach these components together. Examination

revealed that similar-looking substances were found in other seatbacks throughout the

airplane, many of which did not exhibit thermal and/or significant mechanical damage.

According to Weber,217 the location and appearance of the red/brown-shaded substance

was consistent with an adhesive used in the construction of the seats. Subsequent

laboratory testing at NASA of several samples of the substance taken from the seatbacks

identified these substances as being consistent with adhesives.218

1.16.4.8 Examination of Explosive Residue 

Examination of recovered wreckage revealed trace amounts of explosive residue

on three samples of material from three separate locations in the airplane wreckage. These

material samples were submitted to the FBIís laboratory in Washington, D.C., with many

other material samples for analysis. The pieces on which these traces were found were

described by the FBI as a piece of canvaslike material and two pieces of floor panel;

however, the exact locations of the traces were not documented. According to the FBIís

laboratory report,219 analysis of each of the three material samples revealed that they

contained traces of different explosives: one contained cyclotrimethylenetrinitramine

(RDX), one contained nitroglycerin, and one contained a combination of RDX and

pentaerythritol tetranitrate (PETN). 

Subsequent investigation revealed that the accident airplane had been used for a

dog-training explosive detection exercise at St. Louis-Lambert International Airport

(STL), St. Louis, Missouri, on June 10, 1996. During that exercise, an airport police

officer placed five training aids containing explosives (one of which was cracked and

could have released small amounts of explosives) in various locations inside the passenger

cabin. Training aids were placed in the following locations: (1) in row 1, in the center

armrest compartment separating seats 1 and 2; (2) in a small closet at the rear of the upper

217 Weber is the manufacturer of the seats located in the passenger cabin region located above the CWT,

in zone C. 

218 For further information, see the Factual Report on Red/Brown-Shaded Substance Found on Passenger

Seats, dated October 2000. 

219 See FBI file No. 265A-NY-259028, dated March 4, 1997. 
Factual Information 119 Aircraft Accident Report

deck; (3) in row 10, inside the rear pouch of the seat back in front of seat 9; (4) in row 20,

in the overhead compartment above seats 1 and 2; and (5) beside exit door R3 (see

figure 20 in section 1.6.3.2 for an airplane diagram).220

1.16.4.8.1 FAA Studies on the Effects of Sea Water on Explosive Residue 

Contamination

FAA personnel at the FAAís William J. Hughes Technical Center in Atlantic City,

New Jersey, conducted tests to study the effects of sea water on cloth and aluminum

airplane parts that had been exposed to explosives contamination.221 Pieces of aluminum

from the wing of a commercial airplane and pieces of airplane seat covers were

contaminated with explosive residue by manually deposited explosives and by exposure to

an actual explosion and were then immersed in sea water. In both cases, when the cloth

and aluminum airplane parts were examined after 2 days of immersion in sea water, the

explosive residues were completely dissipated. The FAAís report concluded, ìour

experiments have shown that their [sic] is very little likelihood that blast deposited

explosive materials remain very long on cloth or aluminum aircraft parts after immersion

in [sea water].î The FAA attributed this, in part, to the explosive residues being

ìsomewhat soluble in [sea water].î 

1.16.4.9 Wing Center Section and Center Wing Fuel Tank Structural 

Response and Failure Information 

Boeing performed structural response calculations to determine the failure

pressures and response frequency222 for the beams and spars in the WCS/CWT.223 On the

basis of their calculations, Boeing engineers generated a range of pressure differentials

required to fail these members.224 The failure pressure differentials for the SWBs and

spars of the WCS/CWT are summarized in table 5. The ìminimumî pressure column in

the table corresponds to the ìminimum initial failure strengthî determined by

conventional stress analysis methods used in commercial airplane design to ensure that

minimum strength will always exceed regulatory requirement. The ìmaximumî pressure

column in the table corresponds to the ìestimated maximum initial failure strengthî

determined by using detailed finite element computer models.

220 For additional information, see the Hazardous MaterialsóSecurity Group Chairmanís Factual Report,

dated April 27, 1999. 

221 For additional information, see the FAAís report, ìImmersion Studies of Aircraft Parts Exposed to

Plastic Explosives,î which is appended to the Safety Boardís Hazardous MaterialsóSecurity Group

Chairmanís Factual Report. 

222 Frequency analysis of the structural members indicated that these members would respond statically

to the rates of pressure differences generated by the combustion processes. 

223 Because nobody has tested a WCS to failure, the failure strengths predicted by Boeing were based

entirely on traditional analytic methods and numerical computations. 

224 See Safety Board Memorandum to the TWA docket, ìBoeing Submission,î dated June 5, 2000. In

addition to the assumptions identified in the memorandum, it should be noted that the failure pressures that

were calculated do not account for progressive failures in which the failure pressure for a beam may be

decreased because of failure or damage of a neighboring beam.
Factual Information 120 Aircraft Accident Report

1.16.5 Jet A Fuel Vapor and Fuel Tank Research and Testing

During its investigation of the TWA flight 800 accident, the Safety Board was

hampered by the lack of available pertinent research data regarding the flammability and

other characteristics of Jet A fuel and its behavior in airplane fuel tanks. In an attempt to

develop data that would be useful in the investigation, the Board assembled a team that

included the most renowned experts in the combustion and fuel-related fields and, with the

help of these experts, developed and implemented a research program. The objectives of

this research program were to (1) develop an understanding of the thermal and vapor

environment within the 747 CWT; (2) determine the chemical and physical properties,

flammability, and combustion behavior of Jet A fuel; and (3) develop computer models of

the combustion process within the 747 CWT to assist in the determination of the ignition

location. The researchers performed laboratory experiments, scale-model tests, and

numerical simulations to examine the explosion of Jet A fuel/air vapor under conditions

similar to those that existed within TWA flight 800ís CWT at the time of the accident.

Specifically, the researchers wanted to identify the flammability limits and combustion

behavior of Jet A fuel and the propagation of flames through the compartments within the

CWT structure. Figure 32 shows a flow chart of the fuel-related research conducted to

support this investigation. 

The FAA has certificated commercial jet airplanes based on the assumption that

fuel tanks are considered to always be flammable and has required manufacturers to

demonstrate that any energy inside the tank would be below the MIE for Jet A vapors to

ensure safe operation of the airplane. However, at the time of this accident, specific

information about fuel tank flammability and Jet A thermochemistry was largely

unavailable. The research conducted during this investigation enabled analysis of the

flammability of aircraft fuel tanks and provided a scientific basis for evaluating methods

of reducing fuel tank flammability, as identified in safety recommendations previously

issued to the FAA.225

Table 5. 747-100 wing center section/center wing fuel tank 

SWBs and spars failure strengths.

Component Minimum (psi) Maximum (psi)

Front spar 20 25 to 30

SWB3 20 25

SWB2 20 30 to 35

Mid spar 20 35 to 40

SWB1 25 45 to 50

Rear spar 30 45 to 50

225 For more information about these safety recommendations, see section 1.18.3.
Figure 32. A flow chart of the fuel-related research conducted to support this investigation. Probab

L

Full-scale computations1/4 scale CWT combustion experiments Fuel Chemistry Characterization Early Event Damages in Wreckage Structural Failure Analysis Indicates areas of basic research that were developed for the investigation. Flame-front quenching experiments Flammability of Jet A fuel vapors Fuel weathering Fuel Mass Loading Ignition Energy of Jet A Vapors Ground testing Flight testing Rules-based system analysis Combustion properties of Jet A CWT thermal and vapor environment In-tank vapor sampling Flammability Exposure and Reduction Analysis Safety Board Recommendations to Reduce Flammability Code Development and Validation Accident Investigation of CWT Explosion 
Factual Information 122 Aircraft Accident Report

The research associated with this accident investigation included the use of

standardized tests, engineering models, state-of-the-art scientific techniques, development

of new methods, and extension of existing methods. For example, the research team

adapted numerical tools and developed computational models for use in full-scale

simulations of 747 CWT combustion that explored the effects of various parameters and

assumptions (such as ignition locations within the CWT). Additionally, a computational

rules-based analysis method was developed to compare the simulations against the

observed damage to the recovered wreckage. (The goal of this computational effort was to

narrow down the number of probable ignition source locations within the CWT.)226

Flight and ground tests were performed to determine, as accurately as possible, the

thermal and vapor conditions that were present inside the accident airplaneís CWT at the

time of the accident. These tests, in part, identified specific conditions present in the

accident airplaneís CWT (temperature, altitude [pressure], mass loading,227 and

weathering)228 considered important in identifying the behavior of the combustion that

took place within the accident airplaneís CWT. Using this information, all subsequent

analyses and tests were focused on conditions that best approximated those of the accident

airplane at the time of the accident. 

Analysis of the fundamental chemical and physical properties of Jet A liquid and

vapor, calculation of vapor pressure, and molecular weight produced information that was

then used to predict fuel vapor concentrations in the accident airplaneís CWT at the time

of the accident. On the basis of hundreds of tests, conducted over a wide range of

conditions, the research team was able to identify the conditions (ignition energies, fuel/air

mixtures) under which Jet A fuel combustion could occur and to quantify the behavior

(flame speeds and peak pressures) that would likely result from such combustion. 

Two independent computer models, using different numerical codes (selected to

complement each other), were developed to predict the propagation of combustion in a

scale model and full-scale model of the 747 CWT. Hundreds of computer simulations

were performed to develop and validate the models. The results of the tests conducted

with the models showed the pressure differences that developed across the structural

members that divide the 747 CWT into compartments. These models were then used to

make calculations of Jet A combustion in a full-scale CWT geometry, for many different

combustion scenarios. 

226 The research efforts and their results are discussed further in subsections of this section. Additionally,

the research efforts were extensively documented in a series of reports that were submitted to the Safety

Board and are available in the public docket for this accident investigation. See subsections for specific

report references. 

227 Fuel mass loading refers to the measure of the amount of fuel relative to the entire volume of its

container. 

228 Weathering is the change in a liquid fuel chemical composition as a result of exposure to

environmental conditions. An example involves heating and pressure changes to a vented aircraft fuel tank,

where preferential evaporation of the lower molecular weight components of the jet fuel occurs, resulting in

a redistribution of the chemical composition of the remaining liquid fuel. 
Factual Information 123 Aircraft Accident Report

Scale-model experiments were also used to validate the numerical analyses and to

demonstrate the propagation of an explosion in a multicompartment fuel tank. During the

scale-model testing, quenching229 was identified as an issue because of the complex nature

of the CWT. As a result, additional experiments and engineering modeling were

accomplished to address the issue of quenching. 

Ultimately, the results of the computer model research were used in an attempt to

identify the location of probable ignition within the CWT. That effort required the use of

physical evidence, data, and estimates derived from the investigation despite substantial

uncertainties and possible interpretative bias. To address these issues, a quantitative

technique based on statistical reasoning, called a rules-based analysis method, was

developed. 

Before the rules-based analysis could be used to compare the results of the

numerical simulations of the explosion process with the observations of damage to the

airplane and estimates of structural responses within the 747 CWT, it was necessary to

identify the damages that resulted from the initial combustion event. Because damages

observed in the wreckage could have been caused by the initial overpressure event, the

subsequent in-flight breakup, impact with the ocean, and/or fire, it was necessary to

develop a consensus as to which damages resulted from the initial combustion event.

Thus, Safety Board investigators and Boeing engineers identified these initial damages

and assigned levels of confidence to each; these damages and levels of confidence were

then coded into the rules-based analysis method. Additionally, Safety Board investigators

and Boeing engineers estimated the pressures at which various CWT structural members

would fail and assigned levels of certainty to each. (These uncertainties existed because a

CWT had never been tested to failure. As a result, Boeingís estimates provided a range of

estimated failure pressures.) 

Through incorporation of numerical simulations, observed damages, and structural

failure analysis, the rules-based analysis method provided an indication of how consistent

each numerical scenario was with the observable damage to the wreckage and provided

investigators with a means of comparing and evaluating each possible ignition source

location within the CWT. Further, the rules-based analysis resulted in a scientifically

based, quantitative method of analyzing the CWT combustion event in general. 

The following subsections describe the ground and flight tests, vapor sampling, Jet

A fuel and vapor experiments, 1/4- and full-scale model tests, computer modeling, and the

rules-based analysis method in greater detail. Additionally, sections 1.16.5.6 and 1.16.5.7

describe the full-scale 747 CWT explosion test conducted in Leicestershire, England, and

the Boeing/USAF E-4B CWT fuel heating study, respectively. 

229 Quenching is the extinguishment of a combustion flamefront; often as a result of decreased

temperature or propagation through a passageway, such as an orifice or a vent. 
Factual Information 124 Aircraft Accident Report

1.16.5.1 Testing Conducted to Identify the Conditions that Existed in the 

Accident Airplane at the Time of the Accident 

1.16.5.1.1 747 Flight Tests at JFK

In mid-July 1997, the Safety Board conducted a series of flight tests230 to obtain

data regarding the operating temperatures, pressures, and accelerations within and around

a 747-100 series airplaneís CWT under conditions similar to those that existed on TWA

flight 800. The airplane used for these flight tests was a 747-121 leased from Evergreen

Airlines.231

The Safety Board determined the type and location of additional sensors and

instrumentation to be installed and developed the flight test procedures for its series of

flight tests, with input from parties to the investigation. Additional instrumentation was

installed in the test airplane (by Boeing)232 to measure the following: CWT ullage and fuel

temperatures at 48 locations,233 internal and external CWT surface temperatures at

42 locations, air conditioning pack bay temperatures at 20 locations (beneath the CWT),

air conditioning pack surface temperatures at 13 locations,234 vibration measurements of

the CWT lower skin panel,235 and EMI data. In addition, temperature sensors were

installed in the No. 3 fuel tank at each vent and near the high point of each fuel pump

power conduit, on the inside surface of the top wing skin in the No. 3 fuel tank vent

230 In total, nine flight tests were conducted. Six flight tests were performed by the Safety Board and are

documented in the Flight Test Group Chairmanís Factual Report, dated November 19, 1997. The remaining

three flight tests were performed by Boeing and FAA personnel. The purpose of these three flight tests was

to examine the effects of certain hardware modifications on heating in fuel tanks; the data from these flights

are considered proprietary. 

231 The 747-100 series included the -121 variant used in these flight tests and the -131 variant (the

accident airplane). According to Boeing, the fuel systems in these two 747 variants are the same, except for

the method of vent stringer flame protectionóthe accident airplane had an STP system (which was

developed by Boeing for TWA after a TWA 707 fuel tank explosion in Rome, Italy), whereas the flight test

airplane used a flame suppression canister. The 747-121 test airplane was flown by Boeing pilots under an

experimental airworthiness certificate during the flight tests. In addition to the Boeing flight crew, the Safety

Boardís Flight Test Group Chairman and Program Test Director and Boeingís Test Director, Flight Analysis

Engineer, and Flight Instrumentation Engineer were on board the airplane for all flight tests. For those tests

during which vapor sampling was accomplished, a vapor sampling operator was also on board the airplane.

Evergreen Airlines staff were used for mechanical dispatch, maintenance, operations, and ground support

for the flight tests. 

232 For additional information regarding instrumentation, see the Flight Test Group Chairmanís reports

(Exhibit 23B), ìFlight Test Planî and (Exhibit 23E) ìTest Item Requirements List (TIRL), Instrumentation

Locations and Flight Test Schedule,î dated November 19, 1997. 

233 Each fuel bay was instrumented individually to identify temperature variations throughout the CWT. 

234 Temperature sensors were placed in the following locations on each air conditioning pack: on the

engine bleed air duct near the flow control valve, at the water separator outlet, on the top of the heat

exchanger inlet, and on the top of the compressor outlet. In addition, a temperature sensor was placed on the

side support housing and on the heat exchanger exhaust louver. 

235 It was postulated that the CWT might have been exposed to sufficient vibration to result in very small

droplets of the remaining CWT fuel becoming airborne (known as lofting), resulting in a more flammable

fuel/air mixture in the CWT. Examination of the vibration measurement data revealed that the CWT lower

skin panel was not subject to vibration of the energy or frequency that would be necessary to produce fuel

lofting. 
Factual Information 125 Aircraft Accident Report

stringer, and in the surge tanks (at four locations in the right surge tank and at three

locations in the left surge tank). 

Evergreen Airlines provided ballast weight and arranged fueling such that the

airplaneís weight and balance, dispatch fuel load, and takeoff fuel load matched those of

TWA flight 800, as closely as possible, for test flights in which duplication of conditions

was required. The Jet A fuel loaded in the CWT during the flight tests was originally

loaded on an Olympic Airways 747 in Athens, Greece, and flown to JFK on a regular

service flight, similar to the flight flown by the accident airplane before the accident flight.

The fuel was purchased from Olympic Airways, unloaded from the Olympic Airways 747

CWT into a fuel truck, then loaded into the flight test airplane CWT before the first test

flight. Six Safety Board flight tests were conducted with multiple variations in aircraft

operations. One of these six tests, the TWA flight 800 emulation flight test, was designed

to replicate the configuration, ground operation, and flight profile of the accident airplane

as closely as possible. The following is an explanation of this flight test:236

ï TWA Flight 800 Emulation Flight TestóDuring this flight test, the pilots

attempted to duplicate TWA flight 800ís preflight actions/operations, takeoff,

and ascent as closely as possible. The flight test airplane took off within

1 minute of the time of day that the accident airplane took off, 2 days before

the 1-year anniversary of the accident flight. Slight variations from the TWA

flight 800 climb profile were necessary to comply with ATC instructions;

however, the overall climb profile matched that of the accident airplane within

1,000 feet, elapsed times within 1 minute, and airspeeds within 20 knots

throughout the climb. 

ï The flight test airplaneís CWT contained 50 gallons of Jet A fuel (obtained

from the Olympic Airways 747 from Athens, Greece),237 and the main wing

fuel tanks were refueled at JFK to approximate TWA flight 800ís fuel loads.

During the refueling, the outside air temperature (measured on the tarmac near

the airplane) was 88° F.238 The temperature of the fuel loaded on the airplane

ranged from 88° to 91° F. Air conditioning pack Nos. 1 and 3 were operated at

the ìfull-coldî setting for about 2 3/4 hours before the flight test pilots began to

start the airplaneís engines. The airplane began to taxi for takeoff about 1957

and took off about 2021. The air conditioning packs were turned off for

takeoff, and all three air conditioning packs were started after takeoff (in

accordance with TWA procedures) and remained operational for the remainder

of the flight. As the airplane climbed through 10,000 feet msl, the pilots

initiated a crossfeed procedure outlined in the TWA Flight Handbook for fuel

management. During the next 5 minutes, fuel tank No. 3 provided fuel to the

236 As previously stated, additional information regarding the other five flight tests is included in the

Flight Test Group Chairmanís Factual Report, dated November 19, 1997. 

237 Fuel from the Olympic Airways 747 CWT was used so that vapor samples taken from the test

airplaneís CWT during the flight tests would more closely represent the weathered fuel in the accident

airplaneís CWT. This fuel was tested after the flight test; its flash point was determined to be 116° F. 

238 Meteorological records indicate that the temperature at JFK when the accident airplane departed was

82° F. 
Factual Information 126 Aircraft Accident Report

Nos. 1, 2, and 3 engines with both No. 3 boost pumps operating, while fuel

tank No. 4 provided fuel to the No. 4 engine with both No. 4 boost pumps

operating. For the remainder of the flight, the pilots performed the fuel

management procedures described in the TWA Flight Handbook. The pilots

flew the ascent profile recorded by the accident airplaneís FDR (as closely as

possible) until the airplane passed through 13,700 feet msl, then continued to

climb at a constant rate. The airplane leveled off at 19,000 feet msl and

remained there for about 2 hours before it returned to JFK about 2241. Three

CWT fuel/air vapor samples were taken during this flight test (during taxi for

takeoff and as the airplane climbed through 10,400 and 14,200 feet msl). 

Examination of the temperature data collected during the emulation flight test

indicated that the highest ullage temperature measured within the CWT was 145° F and

that it occurred in the left mid bay just before the airplane began to taxi for takeoff.

Examination of the temperature data also indicated that the highest ullage temperature

measured at 13,700 feet msl was 127° F and that it occurred in the left mid bay (see

figure 33). Table 6 shows the maximum and minimum recorded ullage temperatures at

13,700 feet msl during the emulation flight test.239

The Safety Board also measured the ambient air and component surface

temperatures within the air conditioning pack bay. At the time that the flight test airplane

was pushed back from the gate (after operating on the ground for 2 3/4 hours with the

Nos. 1 and 3240 air conditioning packs operating), air conditioning pack component

surface temperatures ranged from 250° to 350° F, and ambient air temperatures within the

pack bay ranged from 148° to 228° F.241

239 Temperature sensors were located in the middle of each bay at three height locations: 3 inches below

the top of the CWT (upper CWT), 3 inches above the bottom of the CWT (lower CWT), and halfway

between the top and bottom of the CWT. Bays 1 and 2 also had temperature sensors located 3 inches in from

the left sidewall and 3 inches in from the right sidewall, at the same height locations. 

240 These are the farthest forward air conditioning packs, mounted just aft of the front spar on either side

of the keel beam. 

241 These temperature ranges reflect variations in the temperature sensorís proximity to the operating air

conditioning components. For example, the temperatures measured near the bleed air inlets to air

conditioning pack Nos. 1 and 3 were at the high end of the range, whereas temperatures measured near the

bleed air inlets to air conditioning pack No. 2 were at the lower end of the range, and the temperatures

measured at the compressor outlets for air conditioning pack Nos. 1 and 3 were in the low- to

mid-temperature range. After the airplane became airborne, these temperatures decreased, but not

significantly enough to rapidly change the temperatures within the CWT. 
Factual Information 127 Aircraft Accident Report

Figure 33. Side view, top view, and three-dimensional view of wing center section/center 

wing fuel tank bays.

Dry Bay

Bay 1

Bay 2

Left Mid Bay Right Mid Bay

Left Aft Bay Right Aft Bay

Side View

Top View

3D View

Forward

Forward

Forward
Factual Information 128 Aircraft Accident Report

As previously discussed, the CWT is configured within the airframe such that the

air conditioning packs are located in an enclosed bay, directly below and very near242 the

bottom surface of the CWT.243 Because of the high temperatures at which the air

conditioning packs operate, the ambient air temperature within the pack bay increases

when the air conditioning packs are operated. As a result of these elevated temperatures,

heat in the pack bay is transferred into the CWT through its bottom surface. Some of this

heat is transferred out of the CWT (for example, into the wing main fuel tanks or the main

landing gear wheel well), and some of it remains in the CWT. This heat flow results in a

significant increase in the temperature of the CWT fuel and ullage, which was shown in

table 6. The temperature variations within the CWT are illustrative of this heat flow. In

general, the hotter temperatures in the CWT ullage were located in the lower, center

portions of the CWT (nearest the heat source), whereas cooler temperatures were

measured around the sides and top of the CWT (where heat was transferred out of the

CWT). 

Examination of the temperatures measured in the surge tanks, vent stringers, and

the No. 3 fuel tank during the TWA flight 800 emulation flight test indicated that these

temperatures were about the same or slightly warmer than the outside air temperature

during the ground portion of the flight test (airplane temperatures of 85° to 90° F versus

Table 6. A summary of the maximum and minimum recorded ullage temperatures at 

13,700 feet msl during the emulation flight test.

Ullage Bay Location

Maximum Temperature in ° F 

(including temperature 

sensor height location)

Minimum Temperature in ° F 

(including temperature 

sensor height location)

Bay 1a

óCenter 117 (lower CWT) 114 (upper CWT)

Bay 1óLeft side 111 (lower CWT) 107 (lower CWT)

Bay 1óRight side 110 (lower CWT) 101 (upper CWT)

Bay 2b

óCenter 120 (lower CWT) 117 (upper CWT)

Bay 2óLeft side 110 (lower CWT) 102 (upper CWT)

Bay 2óRight side 109 (lower CWT) 102 (upper CWT)

Left Mid Bay 127 (lower CWT) 114 (upper CWT)

Right Mid Bay 114 (lower CWT) 106 (upper CWT)

Left Aft Bay 120 (lower CWT) 113 (upper CWT)

Right Aft Bay 112 (lower CWT) 103 (upper CWT)

a. In this section, the term ìbay 1î refers to the fuel bay between SWB3 and SWB2. 

b. In this section, the term ìbay 2î refers to the fuel bay between SWB2 and the mid spar. 

242 The air conditioning packsí distance from the bottom of the CWT varies from about 2 to about

12 inches. See figure 7 for a side-view diagram of the air conditioning packs/CWT. 

243 The air conditioning packs are also located under the CWT in other Boeing and Airbus airplanes;

however, in some Airbus airplanes, the air conditioning pack bays are ventilated. In contrast, in the DC-10,

MD-11, and L-1011 series airplanes, the air conditioning packs are located in the nose, and in the Fokker

F-100 and MD-80 airplanes, the air conditioning packs are located in the tail. 
Factual Information 129 Aircraft Accident Report

outside air temperatures of 82° to 88° F [depending on the time and location of the

temperature reading]). As the test airplaneís altitude increased after takeoff, the

temperatures measured in the surge tanks, vent stringers, and No. 3 fuel tank decreased.

When the TWA flight 800 emulation flight test climbed through 13,800 feet, the

temperatures in the surge tank and vent stringers had decreased to 68° F, whereas the

temperatures measured in the No. 3 fuel tank had decreased to 65° to 75° F (depending on

the temperature sensor location).244

1.16.5.1.1.1 Fuel Vapor Sampling and Analysis from 747 Flight Tests at 

JFKóDesert Research Institute 

Personnel from the Desert Research Institute (DRI), Boeing, and the Safety Board

designed a vapor sampling system, which was installed and functionally tested on the

flight test airplane before the first test flight.245 The sampling system consisted of

preevacuated, 1-liter stainless steel sampling bottles,246 which were plumbed through a

central manifold and mounted in the airplaneís forward cargo compartment. A stainless

steel sampling tube extended from the manifold in the forward cargo compartment

through the front spar and an access panel on SWB3. This sampling tube had an outside

diameter of 1/8 inch, was 25 feet long, and terminated with an open end located about

12 inches aft of SWB3 within bay 2 (about 30 inches above the bottom of the tank and

35 inches left of the CWT centerline), where the fuel vapor samples were drawn. Vapor

samples were collected during taxi and as the airplane climbed through about 10,000 and

14,000 feet msl during the emulation flight test. The vapor sample bottles were

subsequently transported to DRI, where scientists measured the vapor samples for

hydrocarbon fuel concentrations. 

Review of the data obtained from the emulation flight test revealed that the CWT

vapor samples had fuel/air mass ratios247 of 0.034, 0.046, and 0.054 on the ground during

taxi and as the airplane climbed through about 10,000 and 14,000 feet msl (see table 7).

According to scientific literature248 and tests conducted by experts at the California

Institute of Technology (CIT), the lower flammability limit (LFL) of Jet A fuel is at a

fuel/air mass ratio of 0.032 to 0.038.249

244 For additional information regarding the flight test results, see Flight Test Results: TWA 800

Emulation Flight (Exhibit 23F). 

245 For additional information regarding the vapor sampling mechanism and the handling and testing of

the vapor samples collected, see the Flight Test Group Chairmanís Factual Report, dated November 19,

1997. Also, see the DRI report, ìSampling and Analysis of Vapors from the Center Wing Tank of a Test

Boeing 747-100 Aircraft,î dated November 1997. 

246 Two sampling bottles were used for each vapor sample acquisition. One bottle was used to purge the

sampling line and manifold, and the other, subsequently, was used to collect a valid vapor sample.

247 The fuel/air mass ratio is a measure of the mass of fuel in a mixture divided by the mass of the air in

that mixture. In the context of this flight test sample, it is a measure of the mass of the fuel in the ullage

divided by the mass of the air in the ullage. 

248 Nestor, L. 1967. Investigation of Turbine Fuel Flammability within Aircraft Fuel Tanks. Final Report

DS-67-7. Naval Air Propulsion Test Center, Naval Base, Philadelphia, Pennsylvania. Also, Shepherd, J. E.;

Nuyt, C. D.; and Lee, J. J. 2000. Flashpoint and Chemical Composition of Aviation Kerosene (Jet A).

Graduate Aeronautical Laboratories, CIT. Explosion Dynamics Report FM99-4. 
Factual Information 130 Aircraft Accident Report

During subsequent flight tests, investigators collected additional vapor samples

from the CWT ullage. The fuel in the CWT was not consumed, nor was the CWT refueled

during the series of flight tests in which vapor samples were collected. Examination of the

series of vapor samples showed that the vapor composition gradually changed through a

process known as fuel weathering. Subsequent flash point testing (conducted at CIT)250 of

liquid fuel samples collected at each flight test did not indicate a significant change in

flash point (less than 3° F) as the fuel weathered during the series of flight tests. 

1.16.5.1.2 Testing of Fuel Samples from TWA Flight 881 (Arriving at JFK 

from Athens, Greece) 

On October 1, 1996, a fuel sample (about 50 milliliters) was obtained from the

CWT of a TWA 747 that had just arrived at JFK from Athens, Greece, as TWA flight 881.

Because TWA flight 881ís CWT had been refueled in Athens and was believed to have

encountered conditions (flight profile, thermal environment, and weathering) similar to

those encountered by the accident airplane during its flight from Athens to JFK (before it

departed as TWA flight 800), the Safety Board considered this fuel sample to be

representative of the fuel in the accident airplaneís CWT when the overpressure event

occurred. This fuel sample was taken to Saybolt Internationalís laboratories for

standardized fuel testing. The flash point of this fuel was measured at 114° F, and it was

noted that the sample contained an antistatic additive that resulted in the sample having 90

picosiemens/meter electrical conductivity.251

249 From May 29 to 31, 1998, Boeing performed additional ground tests, with technical support from the

Safety Board, to investigate the 747-100 CWT thermal and vapor environment. These tests were conducted

at Pinal Air Park in Marana, Arizona, using a 747-100 series airplane leased from Evergreen Airlines and

were similar in scope to the JFK flight tests conducted by the Safety Board. Testing was performed in a hot

ambient environment (outside air temperatures above 95° F), with all three of the airplaneís air conditioning

packs operating. The tests were conducted in the following conditions: (1) a nearly empty fuel load in the

CWT (50 gallons, similar to TWA 800), with uninsulated air conditioning packs; (2) 50 gallons of fuel in the

CWT, with thermally insulated air conditioning packs; and (3) 12,000 pounds of fuel in the CWT, with

uninsulated air conditioning packs. Temperature and vapor measurements of the CWT indicated that a

flammable condition existed during ground operations during each of the three tests. For additional

information, see Summary Data Report: B-747-100 Center Wing Tank Ground Testing at Marana, Arizona,

dated January 20, 2000. 

Table 7. Fuel/air mass ratios and ullage temperatures measured at the

bay 2 sampling port during the TWA flight emulation flight test.

Sample

Temperature of Ullage at 

Sample Location in ° F Fuel/air Mass Ratio

Emulation flight test, taxi 123 0.034

Emulation flight test, about 

10,000 feet msl 115 0.046

Emulation flight test, about 

14,000 feet msl 117 0.054

250 Shepherd, Nuyt, and Lee.
Factual Information 131 Aircraft Accident Report

1.16.5.2 Laboratory Testing of Jet A FuelsóCalifornia Institute of 

Technology

A review of Jet A combustion data available when the TWA flight 800 accident

occurred revealed that insufficient research data were available to adequately address the

specific issues germane to the characteristics of a CWT explosion.252 Therefore, the Safety

Board contracted with explosion dynamics experts at the CIT Graduate Aerodynamics

Laboratory to conduct a comprehensive experimental and analytical investigation of the

combustion behavior of Jet A fuel. This included flammability research to identify

conditions under which Jet A fuel/air mixtures are flammable and combustion research to

determine the characteristics (for example, explosion peak pressures and flame speeds) of

Jet A fuel combustion. The issues of ignition energy, temperature, altitude (pressure),

mass loading, weathering, and fuel chemistry253 were investigated because of their

criticality to the explosion of the CWT. 

The flammability research focused on defining the conditions at which Jet A fuel

vapor was found to be flammable.254 A precision test fixture was developed to generate

fuel vapors over a large range of temperature and pressure conditions, and an electronic

spark system was developed to ignite the vapors while accurately measuring the ignition

energy supplied. Hundreds of tests were performed255 to determine the flammability limits

of the fuel vapor, as a function of ignition energy, fuel temperature, pressure (to simulate

altitude), fuel mass loading, and fuel weathering. The test conditions were selected to

251 Electrical conductivity for fuel is the ability of the fuel to dissipate a static charge, and the unit of

measure is picosiemens per meter. For additional information regarding the results of fuel tests conducted

during this investigation, see Powerplants Group Chairmanís Factual Report, dated February 25, 1997. This

fuel sample was also analyzed for chemical composition. For additional information, see the University of

Nevada report, ìSampling and Analysis of Vapors from the Center Tank of a Test Boeing 747-100 Aircraft,î

dated November 1997. 

252 According to the experts at CIT, the Jet A flammability data at the time of the accident included

limited data on flammability properties found in the industry standard references (Kuchta, J. M. et. al. 1985.

Aircraft Mishap Fire Pattern Investigations. Final Report APWAL-TR-85-2057. Aero Propulsion

Laboratory.; and Zabetakis, M. G. 1965. Flammability Characteristics of Combustible Gases and Vapors.

Bulletin 627. Bureau of Mines.), much of which was without direct attribution. However, the CIT experts

indicated that much of the Jet A flammability data that existed at the time of the accident appeared to have

been derived from the following three technical reports: Nestor; Ott, E. 1970. Effects of Fuel Slosh and

Vibration on the Flammability Hazards of Hydrocarbon Turbine Fuels within Aircraft Fuel Tanks. Technical

Report AFAPL-TR-70-65. Fire Protection Branch of the Fuels and Lubrication Division. Wright Patterson

Air Force Base, Ohio; and Kosvic et. al. 1971. Analysis of Aircraft Fuel Tank Fire and Explosion Hazards.

Technical Report AFAPL-TR-71-07. Air Force Aero Propulsion Laboratory, Wright Patterson Air Force

Base, Ohio. 

253 Fuel chemistry variations commonly result from different formulations of Jet A fuel and can be

affected by variations in fuel supply, fuel supplier, fuel handling, and/or mass loading. 

254 For additional information regarding these tests, see CIT/Graduate Aeronautical Laboratory. 1999.

Spark Ignition Energy Measurements in Jet A. Explosion Dynamics Laboratory Report FM97-9. Also, see

CIT/Graduate Aeronautical Laboratory. 2000. Spark Ignition Measurements in Jet A: Part II. Explosion

Dynamics Laboratory Report FM99-7. 

255 Initially, these tests were performed using single component hydrocarbon fuels (methane and

propane) to validate the experimental procedures and test results when compared to previously published

research on these fuels. However, subsequent tests were performed using Jet A fuel/fuel vapors. 
Factual Information 132 Aircraft Accident Report

thoroughly document Jet A fuel vaporsí characteristics under conditions that were

determined to have existed in the accident airplane. 

The explosion dynamics experts at CIT also conducted research to determine the

combustion characteristics of Jet A fuel and how those characteristics are affected by

different fuel temperatures, pressures (simulating different altitudes), fuel mass loading,

fuel batch source, and fuel weathering.256 Combustion characterization required hundreds

of experiments257 of the ignition of Jet A fuel vapors in various size test vessels.

Measurements of pressure histories and explosion peak pressures were made during tests

conducted under a multitude of conditions. Flame speeds were calculated over the range

of conditions. The flame speeds and explosion peak pressures were compared to

theoretical calculations under ideal conditions. The resultant information allowed

researchers to make more accurate theoretical calculations of the combustion of Jet A fuel

under nonidealized conditions. Additionally, this information was critical to the

development of computer models of Jet A combustion within the CWT.258

Upon completion of the tests and research, the experts evaluated the applicability

of the data acquired in relation to the TWA flight 800 accident conditions. The conclusions

specific to the accident conditions were based on the examination of a similar flash point

Jet A fuel, at a temperature range of 104° to 122° F at 13,800 feet msl, with a fuel mass

loading equivalent to 50 gallons in the CWT. The explosion dynamics experts from CIT

and the Safety Board determined the following:

ï The flammability limits of Jet A fuel are variable and depend (at least) on

ignition energy level, temperature, pressure, and mass loading. The magnitude

of the ignition energy of the fuel vapor for the accident airplaneís conditions

(50 gallons of Jet A fuel in the CWT at a pressure equivalent to 13,800 feet

msl) is estimated to vary from 0.5 J at 104° F to less than 0.5 mJ at 122° F. 

ï At 13,800 feet msl, with fuel mass loading conditions simulating those of the

accident airplane, Jet A fuel259 vapors could be ignited at temperatures as low

as 96.4° F. 

ï Published research involving pure hydrocarbon fuels260 estimated the LFL of

these fuels at sea level to be a fuel/air mass ratio of 0.036 to 0.041. Published

research involving Jet A fuels261 estimated the LFL for Jet A to be a fuel/air

256 For additional information regarding this research, see Jet A ExplosionsóExperiments: Laboratory

Testing, dated November 21, 1997. 

257 Similar to the previously discussed flammability research, initial tests were performed using single

component hydrocarbon fuels for validation purposes, with subsequent tests performed using Jet A fuel/fuel

vapors. 

258 For more information, see section 1.16.5.5.

259 The fuel used in this test was obtained from the Los Angeles International Airport (LAX), Los

Angeles, California, and was determined to have a flash point of 116.6° F. As previously indicated, on the

basis of tests conducted on fuel obtained from the same fuel source in Athens, Greece, the fuel on board the

accident airplane was determined to have a flash point of 114° F. 

260 Kuchta. (The data were taken at sea level.) 

261 Nestor. 
Factual Information 133 Aircraft Accident Report

mass ratio of 0.032 to 0.035, for pressure between 0.4 and 1.0 atmospheres.

Although the LFL was not explicitly tested for during CITís research, the

lowest fuel/air mass ratio ignited was calculated at 0.038 at 13,800 feet msl,

using 80 J spark energy. 

ï Most existing Jet A vapor pressure data are highly contradictory and based on

correlations rather than measurements. Vapor pressure varies between fuel

batches, and measurements can easily be contaminated by dissolved air. The

vapor pressure measurements obtained in the CIT research and the fuel vapor

characterization made by Woodrow262 represent a significant portion of the

available Jet A vapor pressure and chemical composition data. 

ï Simulating the altitude and fuel mass loading conditions of TWA flight 800 in

a laboratory test cell, explosion peak pressures of Jet A vapors were measured

at 39.2 psi (at 104° F) and 52.2 psi (at 122° F) above ambient pressure at

13,800 feet (8.6 psi). 

ï Ignition of Jet A vapors (at sea level) can occur at temperatures significantly

below their flash point temperatures. A survey of Jet A fuels used in the current

research indicated that fuel/air mass ratios calculated at the flash point

temperature range from 0.0354 to 0.0488, which indicates that the flash point

temperature of the fuel would be higher than its temperature at the LFL.

Therefore, flash point temperatures are not a reliable guide for assessing the

explosion hazards of Jet A fuels. Flammability tests in vessels or tubes with the

appropriate mass loading factor, ignition source, test and analysis procedures,

and instrumentation must be used to determine flammability limits. 

1.16.5.3 Jet Fuel Vapor ChemistryóUniversity of Nevada

The Safety Board contracted with the University of Nevadaís Center for

Environmental Sciences and Engineering to analyze the properties of Jet A fuel vapor.263

The objective of these analyses was to characterize fuel vapor chemistry and determine

the vapor concentrations (fuel/air mass ratios) generated by jet fuels under various

conditions. Many different sources of Jet A fuel were tested over a range of temperatures

and fuel mass loadings, encompassing the conditions that existed when the TWA

flight 800 accident occurred. These analyses were also able to quantify the effects of fuel

weathering and altitude on fuel vapor concentrations. One analysis was made of Jet A fuel

obtained from Athens, Greece, shortly after the TWA flight 800 accident.264 This analysis

showed that for the temperature range (104° to 122° F) and fuel mass loading in the

accident airplaneís CWT at the time of the accident, the fuel/air mass ratios were between

0.036 and 0.066. 

262 Woodrow, J. and Seiber, J. 1997. The Laboratory Characterization of Jet Fuel under Simulated Flight

Conditions. Center for Environmental Sciences and Engineering. University of Nevada. Reno, Nevada. 

263 Woodrow and Seiber. 

264 As stated previously in section 1.16.5.1.2, this fuel came from TWA flight 881 and was believed to

most closely represent the fuel within the accident airplaneís CWT at the time of the accident.
Factual Information 134 Aircraft Accident Report

1.16.5.4 Quarter-Scale Center Wing Fuel Tank Model Combustion 

ExperimentsóCalifornia Institute of Technology 

The Jet A explosion tests discussed in section 1.16.5.2 were conducted in

laboratory explosion chambers. The Safety Board determined that it also needed to

investigate the phenomena associated with flame propagation in multicompartment,

interconnected, and vented tanks representative of the accident airplaneís CWT. The

Board contracted with CIT and Applied Research Associates to conduct a comprehensive

series of combustion tests using a 1/4-scale model of the CWT. The objectives for these

tests were to (1) provide a definitive set of combustion data for the development and

validation of computer code models265 of the combustion process within the CWT and (2)

demonstrate the combustion of Jet A fuel vapor at the temperature, pressure (altitude), and

fuel mass loading present in the accident airplane at the time of the accident in a simplified

yet representative CWT model. 

Accordingly, the explosion dynamics experts at CIT designed an experimental test

program in which a CWT model was constructed with length, width, and height

measurements that were roughly one-fourth the size of a full-size 747-100 CWT.266 (The

1/4-scale CWT model was 5 feet long and wide and 1 1/2 feet high.) The 1/4-scale model

was constructed with a steel top and bottom, removable steel partitions, and transparent

high-strength plastic (Lexan) sides, which enabled observation of the combustion event.

Later modifications to the model included temperature and pressure controls, which

allowed researchers to simulate the temperature and altitude of the accident airplaneís

CWT.267 The model was instrumented with pressure and temperature gauges, and an

external camera system was incorporated to view the combustion event. 

Seventy-two tests were conducted, in three separate test phases, over a 2-year

period. The initial phase of testing was conducted using a simulant fuel.268 The second and

third phases of testing were conducted using Jet A fuel. Variations in model

configurations, ignition locations, and fuel vapor conditions were examined. Because the

primary intent of these tests was to provide a validation database for computer model

265 For a description of the computational modeling effort, see section 1.16.5.5. 

266 This size difference results in the volume of the model tank being one sixty-fourth that of the

full-scale CWT. For additional information, see CIT/Graduate Aeronautical Laboratory. 1997. Jet A

Explosions ñ Field Test Plan, 1/4-Scale Experiments. Explosion Dynamics Laboratory Report FM97-17. 

267 For additional information regarding these tests, see CIT/Graduate Aeronautical Laboratory. 2000.

Results of 1/4-Scale ExperimentsóVapor Simulant and Liquid Jet A Tests. Explosion Dynamics Laboratory

Report FM98-6. Also, see CIT/Graduate Aeronautical Laboratory. 2000. 1/4-Scale Testing, Part IIó

Simulant Repeatability Series, Jet A Vapor and Quenching. Explosion Dynamics Laboratory

Report FM99-7. 

268 The simulant fuel (1.4 percent propane, 7 percent hydrogen, and 91.6 percent air) was developed and

used in this test series to permit the observation of multicompartment fuel vapor explosions at the local

ambient temperature and pressure conditions of the test site (near Denver, Colorado), simulating the

combustion behavior of the Jet A fuel vapor in the accident airplaneís CWT at the conditions that existed at

the time of the accident. Using the simulant fuel greatly reduced the experimental effort required (compared

to using Jet A fuel). Using Jet A fuel during the second and third phases of tests required significant

modifications to the test model to elevate the temperature to between 104° and 122° F at a reduced pressure

equivalent to an altitude of 13,800 feet msl. 
Factual Information 135 Aircraft Accident Report

development (which required that the essential physical phenomena of the CWT

combustion event be represented), researchers did not require experimental duplication of

all of the conditions on board the accident airplane. Important distinctions between the

1/4-scale test model and the full-scale CWT and the limitations of the testing were the

following: 

ï The complexities of the full-scale CWT were simplified in the 1/4-scale model.

The features believed to be most significantóbay volumes and lengths and

vent and passageway areasówere scaled from the full-scale CWT, with some

simplifications. For example, variations in tank height and the finer geometric

details of stiffeners and stringers were not included in the model. 

ï It is not clear if the results of the 1/4-scale model testing can scale directly to

replicate full-scale results. The effects of scaling on some features (such as

flow turbulence and flame quenching) are not well understood. 

ï The effects of temperature and fuel vapor concentration variations within a

single bay and between bays could not be examined in this limited

experimental program.269

ï Other simplifications of the 1/4-scale model included the even distribution of

fuel between the bays and a smooth floor geometry. Thus, the role of liquid

layer participation may not be accurately demonstrated. 

ï The testing did not attempt to simulate partition failures, representative of fullscale CWT structural dynamic behavior during the combustion event, and the

effects of such failures upon the combustion dynamics. 

Because of the differences between the 1/4-scale test model and the full-scale

CWT, the interpretations that could be made by directly comparing these experimental

results to a full-scale explosion were limited. Within these limitations, however, analysis

of the experimental results led the researchers to reach the following conclusions

regarding direct application of the 1/4-scale test results to TWA flight 800: 

ï Jet A fuel ignited and combusted during every test using conditions

approximating those that existed in the accident airplaneís CWT at the time of

the accident. 

269 Temperature and vapor concentration variations within a single bay and between bays were observed

in the JFK flight test results. Subsequent ground tests were conducted in Marana, Arizona, in which CWT

temperatures and vapor concentrations were measured with higher resolutions than those measured during

the JFK flight tests. These tests confirmed the existence of temperature and vapor concentration variations

within a single bay and between bays. For additional information, see Summary Data Report: B-747-100

Center Wing Tank Ground Testing at Marana, Arizona, dated January 20, 2000, and the Flight Test Group

Chairmanís Factual Report, dated November 19, 1997. 
Factual Information 136 Aircraft Accident Report

ï The ignition of Jet A fuel in one bay of the 1/4-scale model resulted in

transmission of the flame through the bay passageways and vent stringers and

ignition in neighboring bays, illustrating the behavior of multicompartment

flame propagation.270 Flamefront quenching was also observed to be a

characteristic of flame propagation. 

ï After the combustion flamefront propagated from the ignition bay to a

neighboring bay, the combustion process dramatically accelerated, allowing

explosion pressures to increase rapidly. 

ï In certain tests, pressure levels in bay 1 (the bay between SWB2 and SWB3) of

the 1/4-scale model exceeded those needed to fail SWB3 (as indicated by

Boeingís structural analysis of the full-scale geometry).271

1.16.5.5 Computational Research of Center Wing Fuel Tank Combustion

The Safety Board contracted with two research laboratoriesóSandia National

Laboratories (SNL) and Christian Michelsen Research (CMR)óto develop computer

code models of the combustion process that occurs in a 747 CWT. Both research

laboratories had extensive experience in large-scale explosion modeling and possessed

existing combustion codes that could be applied to the task of modeling the CWT

combustion process. 

This research was intended to develop computer models capable of generating

accurate solutions for different full-scale CWT combustion scenarios.272 Rules-based

analysis,273 which compared the results from each computer solution to the physical

evidence, was then to be applied to these computer solutions to determine the probability

that a particular combustion scenario might have occurred in the accident airplane. The

Safety Board hoped that research involving this combination of computer modeling and

rules-based analysis would indicate the most probable ignition source for the CWT

overpressure event. 

The combustion model development was conducted concurrently by the two

laboratories for 2 years. During this period, model development was continually being

validated using the data generated in the 1/4-scale experimental test program.

Additionally, because each computer model used a different approach toward calculating

the combustion process, the researchers were able to develop and verify their work in

conjunction with each other.274

During the initial development stages, both research laboratories produced

computer models of the 1/4-scale CWT geometry and provided solutions using both the

270 Multicompartment flamefront propagation refers to the progression of a flamefront through the

fuel/air vapor space within and between fuel tank compartments. 

271 For more information, see section 1.16.4.9.

272 Each scenario represents a unique combination of ignition location, ullage temperature, and structural

failure timing, for which a computer solution was calculated. 

273 For additional information regarding the rules-based analysis, see section 1.16.5.5.1. 
Factual Information 137 Aircraft Accident Report

simulant and Jet A fuel. Subsequently, auxiliary computer models and supplemental

experimental tests were conducted to refine the accuracy of certain characteristics of the

Jet A combustion process. The most significant of these involved the issue of modeling

flamefront quenching through compartment passageways. Considerable effort was spent

on this issue, and a simple, limited model was developed to describe this behavior. 

In the final developmental stages, each computer model was revised to model the

full-scale CWT geometry. Two full-scale calculations were performed using the SNL

model, and 32 full-scale calculations were conducted using the CMR model. On the basis

of their analysis of the full-scale computer modeling calculation results, the experts from

both research laboratories concluded the following: 

ï The combustion behaviors of multicompartment flame propagation exhibited

in the full-scale model calculations were consistent with those observed in the

1/4-scale modeling tests. 

ï During some scenarios, the pressures developed in bay 1 exceeded the pressure

required to fail SWB3.275

ï In all of the computer solutions, conditions were calculated that indicated that

quenching could have occurred in some of the vents and passageways of the

full-scale CWT geometry (which has 89 intercompartment passageways

connecting the 6 fuel bays and 2 vent stringers connecting the CWT to the

wing tip vents). 

ï Incorporating the effects of quenching in the calculations appeared to

significantly affect the differential pressure histories that developed across the

internal CWT structural members. This resulted in the solutions having an

enhanced sensitivity to ignition location. 

Because the full-scale computer modeling results indicated that the conditions for

flamefront quenching appeared to exist, and because this behavior affects the development

of compartment pressures, the ability to accurately predict quenching was very important.

However, because the research data regarding the quenching phenomena are limited, a

complete understanding of quenching behavior was not possible, and the issue of

quenching remains unresolved. 

1.16.5.5.1 Study of Computer Model Calculations of Full-Scale Center Wing 

Fuel Tank Combustion

The Safety Board contracted with Combustion Dynamics, Limited (CDL) to

evaluate the consistency between the computer calculations of the full-scale CWT

combustion model and other information and evidence obtained during the investigation.

274 The code used by SNL determines pressure histories from global mass, momentum, and energy

balances. Flame propagation is represented as a moving interface that separates the burned and unburned

gases. The CMR model uses a finite volume technique for solving the mass, momentum, energy, and kinetic

turbulent energy production and dissipation. This model follows the motion of the flamefront, separating

burned and unburned gases. 

275 For more information about the pressures required to fail SWB3, see section 1.16.4.9.
Factual Information 138 Aircraft Accident Report

The Board hoped that by conducting this evaluation for each computer model solution or

scenario (each involving a different combination of ignition location within the CWT,

ullage temperature, and timing of structural failure), it would be possible to narrow the

number of probable ignition location(s) within the CWT. CDL used the rules-based

analysis to evaluate this consistency.276

As previously discussed, the rules-based analysis required the development of the

following areas of the investigation: 

ï identification of damage (or nondamage) observed in the recovered CWT

wreckage that could be attributed to an early event in the breakup sequence,277

such as the failure of SWB3 in the forward direction and the rupture and

forward deflection of the SWB2 manufacturing access door; and 

ï mechanical failure analysis of the CWT structural members, which defined a

range of pressure differentials that could cause failure of each of the CWT

SWBs and spars.278

Using information from these areas, the rules-based analysis developed

probabilities of damage/failure events for the actual (observed in wreckage) and simulated

(calculated by the computer) events. These probabilities were then used to develop a

consistency estimate, which was an indication of how consistent an observed

failure/damage was with the computer modelís calculations for that failure/damage (for

example, pressure differential across that panel). The consistency estimates for each

individual failure/damage event were then combined to produce an overall consistency

estimate, which would indicate the overall agreement between that computer scenarioís

combination of events and the total of the observed damages. 

The rules-based analysis was used to examine each of the 32 scenarios of full-scale

CWT calculations provided by the CMR model. The results showed a positive overall

consistency value between each scenario that was modeled and the observed damages.

Although scenarios involving certain ignition locations produced results with higher

consistency estimates than scenarios involving other ignition locations, none of the

differences in consistency estimates were high enough to permit a determination of the

most probable ignition location with a high degree of confidence. The experts at CDL

attributed this to the uncertainties associated with the full-scale model calculations,279 the

limited amount of airplane damage that could conclusively be considered the result of

early events,280 and the difficulties involved in determining the failure modes and failure

pressure loads of the complex geometries of the CWT structural members (even with full

276 For additional information, see Thibault, P. 2000. Evaluation of Explosion Scenarios. Combustion

Dynamics, Limited. Report CDL-1010. 

277 The early event damage considered in this study/evaluation included damage caused by the CWT

overpressure event and did not include damage caused by the subsequent in-flight breakup of the airplane,

fire, or water impact. 

278 For more information, see section 1.16.4.9 and Safety Board Memorandum to the TWA flight 800

Docket, ìBoeing Submission,î dated June 5, 2000. 

279 For more information about the full-scale model calculations, see section 1.16.5.5.
Factual Information 139 Aircraft Accident Report

utilization of modern computer resources). Therefore, the rules-based analysis did not

provide a definitive determination regarding the probability that any given location within

the CWT was the ignition location. However, the rules-based analysis did reveal that the

pressure differentials produced by an internal fuel/air explosion were consistent with the

overall level of damage observed in the CWT. Further, the experts at CDL indicated that

this conclusion was supported by a simple analysis based on the peak pressures that were

experimentally observed for the combustion of Jet A fuel/air mixtures at the (temperature,

pressure, and mass loading) conditions that existed in TWA flight 800ís CWT at the time

of the accident. 

1.16.5.6 Bruntingthorpe Full-Scale Center Wing Fuel Tank Explosion Tests

Between July 28 and August 2, 1997, the Safety Board, with the assistance of the

British Defense Evaluation and Research Agency (DERA), conducted a CWT fuel/air

explosion test by igniting a propane/air mixture in the CWT of an out-of-service 747

located at Bruntingthorpe Airfield, Leicestershire, England. The Board also detonated

explosive charges at different locations inside and outside of the CWT and documented

the resulting damage.281 Examination of the damage indicated that when metal of the same

type and thickness as the CWT walls was penetrated by a small charge, there was petalling

of the surface on which the charge was placed, pitting on the adjacent surfaces, and visible

hot gas washing damage in the surrounding area. 

Documentation of the damage (direct and collateral) patterns was conducted as a

group activity, with the involvement of technical representatives from several parties to

the investigation, including Boeing and TWA. The tests enabled the Safety Board to

examine the explosion, deformation, and failure processes (crack paths and failure

sequences) produced by the ignition of a fuel/air mixture within the CWT and detonation

of explosive charges. 

It was not possible for the tests (nor was it intended) to exactly reproduce the

explosion of TWA flight 800ís CWT.282 However, weight was added to the test airplane to

simulate some of the loads on the accident airplane as closely as possible. For example,

water was added to the main wing fuel tanks to reproduce the accident airplanesí fuel

loads, the potable water bottles forward of the front spar were filled with water, and

sandbags were placed in some of the passenger seats over the WCS. (The Safety Board

280 Although nearly all of the wreckage has been recovered and, where pertinent, reconstructed, much of

the early event-related damage was altered or affected by the events following the overpressure event,

including the in-flight breakup of the airplane, water impact, fire, and handling during the recovery efforts. 

281 Each of the explosive charges used PE4, a type of plastic explosive using RDX combined with a

plasticizer. In some cases, the charge was coupled with various amounts of bare shot. 

282 For example, it was not possible to simulate the atmospheric pressure, load factors, and/or slipstream

encountered by the accident airplane in flight at nearly 14,000 feet msl nor was it possible to pressurize the

fuselage. As a result of damage from the explosive charge testing (which had taken place previously), the

test airplane exhibited significant damage to the aft fuselage beginning with the STA 1480 bulkhead;

missing sidewall panels, ceilings (including overhead bins), windows, and doors (including the cockpit door

and several main cabin doors); and damage to many other doors (including the forward cargo door) that

precluded closing or latching those doors. However, the airplaneís CWT and wing structure were repaired,

as necessary. 
Factual Information 140 Aircraft Accident Report

notes that, despite these adjustments, the loading of the test airplane was significantly

different from that of the accident airplane. These differences made a direct comparison of

the results of the fuel/air explosion of the CWT in the test airplane with the results of the

fuel/air explosion in the accident airplane of limited use.) 

Portions of the cabin floor structure on the test airplane were modified to

accommodate test hardware and instrumentation.283 The openings in the CWT tank

(including the tank vents to the wings) were hermetically sealed, and a predetermined

amount of propane284 was loaded in the CWT to approximate the amount of fuel vapor in

the accident airplaneís CWT at the time of the explosion. 

The propane/air mixture in the test airplaneís CWT was ignited by detonating an

explosive device located on the right side of the rear spar (about RBL 82). According to

the DERA report, this detonation ìtriggered a major event or events in the [CWT]Ö.a

relatively slow deflagration was evident from the audible output accompanied by a limited

flash, visible externally. Deformation and holing of the fuselage roof was obvious, even

from a distance.î 

Safety Board investigators and representatives from DERA and several parties to

the investigation examined the postexplosion damage to the test airplane and developed a

summary of the failure sequence that they believed showed the progression of the

overpressure resulting from the propagation of the propane/air explosion. The Board

observed that the test parameters used resulted in a significantly more dynamic and

destructive explosion within the test airplaneís CWT than was indicated by the accident

airplaneís wreckage.285 (The catastrophic nature of the damage to the test airplane

indicated that if such an event occurred in flight, it would likely result in the airplane

instantaneously separating into four major components: left wing, right wing, forward

fuselage, and aft fuselage.) 

1.16.5.7 Boeing/U.S. Air Force E-4B Center Wing Fuel Tank Fuel 

Heating Study

In March 1999, the USAF Safety Center, Directorate of Engineering and Technical

Services advised the Safety Board of a USAF E-4B fuel tank heating study that had been

conducted by Boeing for the USAF from 1979 to 1980.286 The E-4B is a military variant

of the commercial 747, manufactured by Boeing, which has a larger CWT fuel capacity

283 Test hardware and instrumentation installed for the tests included pressure, flamefront propagation,

and propane/air concentration sensors throughout the CWT; accelerometers in the passenger seats above the

fuel tank to measure the acceleration forces applied to the seats; a CAM; cockpit and airplane skin

accelerometers; and various pressure gauges in the passenger cabin to obtain complete CVR documentation

of the effects of the explosion. Associated cables and connectors were located within the CWT bays where

necessary. 

284 A propane/air mixture was used because it was very portable and would stay in the vapor state for the

period of time needed for the test, despite the cold climate. 

285 For example, on the test airplane, the force of the explosion shattered SWB3 into a large number of

small pieces, which were propelled forward, whereas on the accident airplane, SWB3 initially only fractured

at its upper chord and then rotated forward into the front spar. 
Factual Information 141 Aircraft Accident Report

than the 747-100.287 The E-4B is also equipped with a 4th air conditioning pack beneath

the CWT. 

According to the Boeing study, the E-4B fuel heating study resulted from USAF

concerns that CWT fuel temperatures could exceed the maximum allowable fuel

temperature (a limit imposed for fuel pump performance, not flammability). Boeingís

study indicated that these concerns were based on the USAFís finding288 CWT

temperatures in excess of the allowable limit. The study also stated that the USAFís

concern regarding the potential problem of high fuel temperatures was supported by

ìreports of a Japan Airlines commercial aircraft operating out of Hawaii that was

experiencing loss of CWT fuel feed during the ascent to altitude.î Boeingís study stated

that the purpose of the study was, in part, ìthe identification of aircraft operational

procedures and hardware modifications which will extend the duration of aircraft

operation before a [CWT] overheat condition occurs.î 

Boeingís study analyzed CWT temperatures during extended periods of ground

operation in extreme (hot, humid) environmental conditions. The study included an

experimental phase, during which fuel temperatures in the CWT were measured, and an

analytic phase, during which computer modeling was used to predict CWT temperatures

under various conditions. CWT fuel temperatures exceeding the allowable limit were

measured during the experimental phase and were predicted under certain operating

conditions by the computer models used in the analytic phase. The study also determined

that although the increase in CWT fuel temperatures was the net effect of several heat

sources and sinks, ìthe major heat load appears to be due to the high temperatures in the

air conditioning equipment [bay].î 

As a result of this experimental and analytic work, the Boeing study proposed

manipulating CWT fuel loads/initial fuel temperatures and ground operation with the air

conditioning pack doors open as operational methods to reduce CWT temperatures. The

Boeing study also concluded that thermal insulation should be installed between the air

conditioning packs and the CWT in the air conditioning pack bay and included

286 The report, ìCenter Wing Tank Fuel Heating Study,î was prepared by Boeing for the USAF, was,

Document No. D226-20582-1. Boeing personnel associated with the TWA flight 800 investigation indicated

that they did not become aware of this study until late in the investigation because the study pertained to the

military variant of the 747 and did not focus directly on flammability. Thus, they were not able to provide

the Safety Board (or the FAA or 747 operators) with the potentially relevant information about 747 CWT

overheating and corrective measures during the early stages of the investigation. According to Boeing, as a

result of this event, the company revised its procedures such that all company reports and resources (military

and/or civilian) are now electronically accessible through key-word searches. Further, Boeing has reported

that its accident investigation procedures have been revised and now include the ability to electronically

search the companyís technical documents. 

287 The E-4B has a CWT capacity of 110,812 pounds (about 17,000 gallons of Jet A at 6.51 pounds per

gallon). This CWT capacity is equivalent to that of a 747-200 (and subsequent 747) model airplanes. The

CWT capacity for the 747-100 is 12,890 gallons. According to Boeing, the increased CWT capacity in the

E-4B/747-200 models was created by using the area between SWB3 and the front spar as a fuel bay, not as a

dry bay, as it was used on 747-100s.

288 For additional information regarding USAF testing, see IOT and E Environmental Testing at Howard

Air Force Base. January 1979.
Factual Information 142 Aircraft Accident Report

preliminary designs for the installation of thermal insulation. Further, the Boeing study

discussed several possible methods that would provide an insulating air layer between the

air conditioning pack bay components and the CWT, including baffling of the air

conditioning pack bay (also termed ìdirectedî ventilation)289 and general forced air

conditioning pack bay ventilation (using either an outside air source or recirculating air

conditioning pack bay air). 

1.16.6 Potential Ignition Source Research and Testing 

During its investigation of the TWA flight 800 accident, the Safety Board

evaluated numerous potential ignition sources of the fuel vapor within the CWT, including

those in which the energy entered the CWT through the FQIS wiring290 and those in which

the energy entered the CWT through other mechanisms. Potential ignition sources in

which the energy may have entered the CWT through the FQIS included radiated

electromagnetic energy from transmitters outside the airplane that coupled to FQIS

wiring; a transfer of electromagnetic force from personal electronic devices (PED) or

other potential emitters within the airplane to the FQIS wiring; a transient voltage/current

spike from wiring associated with other, higher-voltage aircraft systems induced onto

corouted FQIS wiring; or a transfer of voltage from a short circuit from wires carrying

higher voltage to adjacent wires carrying lower voltages, including the FQIS wires (this

short circuit could occur if the insulation covering the conductors is compromised or

through a bridge created by contaminants, such as metal shavings or fluid). 

Other potential ignition mechanisms (those not related to the FQIS) examined by

the Safety Board included a lightning or meteorite strike; a missile fragment; a small

explosive charge placed on or in the CWT; auto ignition or hot surface ignition,291

resulting from elevated temperatures (caused by either engine bleed air leaks, a fire in the

main landing gear wheel well, or a fire in the air conditioning pack bay beneath the

CWT);292 a fire migrating from another fuel tank through the vent (stringer) system; an

uncontained engine failure or a turbine burst in an air conditioning pack; a malfunctioning

jettison/override pump; a malfunctioning CWT scavenge pump; and static electricity. 

289 A concept similar to baffling/directed ventilation is used on Airbus A300, A310, A319, A320, and

A321 airplanes. 

290 According to the FAAís Aging Non-Structural Systems Research Plan, electrical failures can occur in

the following ways: (1) interruption of current in a circuit (also called an open circuit), which the FAA

considered the ìmost benign failure modeÖ[although they] may result in more serious failure in some other

part of the electrical systemî; (2) a bolted short circuit, which will generally trip the airplaneís thermal

circuit breakers and prevent their reset; (3) an intermittent short circuitóthe FAA indicated that ìa high

frequency of intermittent failure may eventually lead to serious consequences either by the more critical

failure of components or by a more critical failure of the interconnect system (e.g., electrical arcing)î; (4) an

intermittent instantaneous discharge (arcing), which the FAA indicated may be ìthe most serious failure

mode for electricalÖcomponentsî; and (5) degraded shielding, which the FAA indicated may result in ìthe

introduction of undesirable noise or electrical energy with potential adverse affects to the safety of the

systems.î 

291 For additional information regarding auto ignition and hot surface ignition, see section 1.16.6.10. 

292 The Safety Board is aware of at least one report of fuel leaking into the air conditioning pack bay in

a 747. 
Factual Information 143 Aircraft Accident Report

1.16.6.1 Arcing Information 

As previously indicated, electrical arcing is defined (in an article published by UL)

as ìa luminous discharge of electricity across an insulating medium.î UL indicated the

following:

The electrical discharge of an arc can involve temperatures of several thousand

degrees Celsius. In determining the heating effects of an arc, the classical Joule

heating equation involving current squared multiplied by resistance (I2R),

however, does not fully explain the heating effects. Although the total power

dissipated in the arc is equal to the total voltage drop in the arc multiplied by the

arc current, power dissipation is not uniform throughout the arc. 

In general, arcing can be divided into two categories: (1) non-contact arcing, and

(2) contact arcing. 

Non-contact arcing is arcing that does not require direct physical contact between

the conductors or ìelectrodesî where the arcing is taking place. Two types of noncontact arcing involving lower voltages are: a) arcing between conductors

separated by insulation that occurs across the surface of the insulation, and

b) arcing between conductors separated by pyrolized (carbonized) insulation.

With arcing between conductors separated by insulation, the mechanism of

initiating an arc between stationary conductors separated by insulation will

depend on the type and geometry of the conductors and insulation between them.

In the case of typical air clearances found in an electrical residential distribution

system, many kilovolts may be required to initiate arcing. 

With arcing between conductors separated by carbonized insulation, also known

as an ìarc-tracking event,î arcing can occur at normal operational voltages. The

resulting fault-current causes the carbon path to open and an arc is established

similar to parting the conductors, as with contact arcing. Carbonized insulation

between opposite polarity conductors or between a line-voltage conductor and

ground can lead to an across-the-line arcing fault or a line-to-ground arcing fault. 

Contact arcing is arcing that involves direct or indirect physical contact between

the conductors, known as electrodes, where the arcing is taking place, such as

arcing between closing or parting conductors making or breaking a circuit. With

this type of arcing, the arc initiation mechanism involves a hot point (essentially

from I2R heating) at the last point of contact when a circuit is being interrupted

(i.e., conductors initially in contact are parting), or at the first point of contact

where a circuit is being established (i.e., conductors that are initially separated and

subsequently come into contact). 

Contact arcing is associated with normal operational arcing that occurs with any

kind of air-gap type electrical switching device. Properly designed switching

devices are capable of withstanding such arcing without excessive contact damage

or generation of excessive heat. Contact arcing may also be associated with arcing

faults due to the unintentional creation or interruption of current. 

1.16.6.2 Wire Short-Circuit Characteristics and Protection 

The contamination observed by investigators on the wiring recovered from the

accident airplane prompted the Safety Board to contract with Lectromechanical Design
Factual Information 144 Aircraft Accident Report

Company (Lectromec) to conduct laboratory research into the short-circuit behaviors of

the following commonly used aircraft wiring: BMS13-42 and -42A Poly-X wires and

Kapton wires.293 The Poly-X wires used for these tests were the most pliable (best case)

samples obtained from a retired 747 (N93117) that was manufactured in 1971 and retired

after 25 years in service, whereas the Kapton wires used for these tests were obtained from

a reel of new wire.294

According to the Lectromec report, ìElectrical Arcing of Aged Aircraft Wiring,î

three types of tests were conducted: wet short-circuit tests in which a 1 percent

saline-water electrolyte solution (NaCl)295 or lavatory fluid was used as a conductor; dry

short-circuit tests in which metal shavings (steel and/or aluminum)296 were used as

conductors; and dry abrasion tests in which metal shavings were placed between

oscillating wires. In all test cases, the test wire bundles were connected to a generator that

provided three-phase, 400-Hz power at 120-volt a.c., in a circuit with 7.5- or 10-amp

circuit breakers and a 1-ohm297 resistor.298 Lectromecís report indicated that during these

tests, three categories of electrical activity were observedóscintillations, flashing, and

strong arcingówhich it described as follows: 

ï Scintillations are high-frequency, micro-discharges that appeared as pinpoints

of light that flickered at the edge of cracks in the wiring insulation.

Scintillations did not result in circuit breakers opening;299 however, during

more lively scintillations, an electrical buzzing or crackling sound could be

heard, and char, or soot, formed on the wire (and adjacent wires) over time.

Lectromec was not equipped to accurately measure the small amounts of

energy in these scintillations.

293 Poly-X wires were tested more extensively than the Kapton wires. For additional information

regarding Kapton and/or Poly-X wiring, see section 1.6.1.3.2 and the Systems Group Chairmanís Factual

ReportóAddendum for Electrical Wiring Information, dated February 10, 2000. Additionally, Poly-X and

Kapton wires are compared in the U.S. Naval Avionics Facility (NAFI) report NAFI-TR-2210, dated

October 19, 1977, which is discussed further in section 1.16.6.3.1. 

294 The types of wires used for individual tests are indicated in the individual test results discussions

and/or tables. As previously discussed, BMS13-42 and -42A Poly-X wires are basically the same except for

the thickness of the external coat (insulation); the BMS13-42 Poly-X used in these tests had an insulation

thickness of about 7 (one thousandth of an inch) mil, whereas the BMS13-42A Poly-X tested had an

insulation thickness of about 11 mil. The Kapton wire used in these tests had an insulation thickness of about

8 to 9 mil. For additional information regarding these types of wires, see Systems Group Chairmanís Factual

ReportñAddendum for Electrical Short Circuit/Arcing of Aged Aircraft Wiring, dated October 7, 1999, and

section 1.6.1.3.2. 

295 These tests were conducted to Society of Automotive Engineers (SAE) test standard AS4373 4.5.9

(Method 509) with variations, such as the use of lavatory fluid as an electrolyte.

296 The metal shavings used in these tests were produced using a 3/16-inch drill bit. The steel shavings

tended to be longer and stronger than the aluminum alloy shavings. 

297 An ohm is the unit of resistance of an electrical conductor, at which the fall of potential is 1 volt when

the current is 1 amp. 

298 Although in this case a resistor was used to simulate the resistance associated with a long run of wire,

typically resistors are used in circuits to limit the amount of current (amperage), and thus power, that can

pass through it. 

299 For additional information regarding circuit breakers, see section 1.16.6.2.1. 
Factual Information 145 Aircraft Accident Report

ï Flashing is an arcing discharge, seen as a single flash of light with an

accompanying popping sound that usually occurred after a test sample had

been scintillating for some time. Rapid, repeated flashing could continue for

several seconds (but not for extended periods of time) and was followed by

periods of dormancy. Typically, 2 to 4 J of electrical energy were transferred

between the source and the target wire, although electrical energy transfers as

high as 90 J were observed. In an April 20, 2000, letter to the Safety Board,

Lectromec stated that the peak power available to the victim wire during a

3/4-millisecond flashing event was 1.3 kilowatts (kW), with a possible energy

dissipation of 400 mJ. During one series of flashing events, Lectromec

measured 8.25 J of energy available in less than 20 milliseconds. Flashing did

not cause the 10-amp circuit breakers protecting the circuit to open,300 but did

result in erosion of insulation and conductors in the wire and adjacent wires

over time.

ï Strong arcing is an arcing discharge that could continue for hundreds of cycles,

typically involving 5 kilojoules (kJ) of electrical energy and that could cause

significant collateral damage to the insulation of adjacent wires. Strong arcing

events often, but not always, resulted in the interruption of power when the

circuit breakers opened (ìtrippedî). Strong arcing was observed in one of three

tests involving a BMS13-42 Poly-X wire that had a relatively thin layer of

insulation (about 7 mil). Tests involving BMS13-42A Poly-X wiring did not

result in the strong arcing that was observed in wet-short testing of Kapton or

BMS13-42 Poly-X.301

According to the report, most of the flashing and strong arcing that occurred varied

widely in intensity and range. In addition, the discharges were observed to be directional

(not a ball-shaped discharge), with sufficient directionality that they missed an adjacent

piece of aluminum in some tests. 

Wet Short-Circuit Tests 

The wet short-circuit tests involved a bundle of seven wires, five of which were

attached to the generator and were considered ìactive,î whereas the other two wires in the

bundle were not attached to the generator and were considered ìpassive.î The wire bundle

was suspended horizontally in a test chamber, with both ends secured to stable platforms.

The insulation on two active wires at the top of the bundle had been intentionally damaged

(cut or sliced) circumferentially, exposing the conductors. The cuts in the two wires were

placed in the bundle such that the cuts were separated by about 6 millimeters (mm). Then,

with the generator providing 120-volt a.c. power, separate tests were conducted with

electrolyte solutions of 1 percent NaCl solution or lavatory waste water dripped on the

wires such that the fluids landed between the cuts in the wire insulation at a rate of 6 to

300 According to a circuit breaker manufacturer, if the current in a circuit rises abruptly (within a few

milliseconds), it can exceed the rated current by more than 1,000 percent before the circuit breaker trips. 

301 Lectromec noted that the strong arcing of BMS13-42 Poly-X was not as energetic as the strong arcing

exhibited by the Kapton.
Factual Information 146 Aircraft Accident Report

10 drops per minute. The experiment was allowed to run for up to 25 minutes or until a

circuit breaker tripped. In some cases circuit breakers were reset and, the test continued.302

Table 8 shows the results of the wet short-circuit tests conducted by Lectromec. 

According to the Lectromec report, during these tests, scintillations were generally

visible soon after the tests were started. During the tests in which flashing occurred,

flashing events frequently repeated at various intervals (ranging from several flashes per

second to more than 1 minute apart) throughout the test. Oscillogram measurements

indicated that the peak electrical energy dissipation (about 90 J, with power peaks between

3 and 6 kW)303 during these flashing events occurred during a test in which a wet-short

was initiated by dripping lavatory fluid onto Poly-X (BMS13-42A) wiring. The estimated

peak electrical energy dissipated before the circuit breakers tripped during the tests in

which strong arcing was observed was about 2 kJ. 

302 The 25-minute limit was selected because it exceeded the length of the accident airplaneís flight; most

of the tests also exceeded the 2 1/2-minute interval between the CVR-recorded comment regarding the

ìcrazy fuel flowî and the explosion. In one undocumented wet short-circuit test, the test continued with

some intense flashing observed for more than 8 hours without tripping a circuit breaker. This test was

eventually terminated at the end of the Lectromec work day. 

Table 8. Results of wet short-circuit tests conducted by Lectromec.

Type of Wire Electrolyte Duration

Qualitative 

Description of 

Visual Effect

No. of Circuit 

Breakers that 

Tripped

Length of 

Damage/Char 

Buildup

Poly-X 

(BMS13-42A)

1 percent 

NaCl 20 minutes Some intense flashing 0 1/2 inch

1 percent 

NaCl 25 minutes Some intense flashing 0 5/8 inch

Lavatory fluid 10 minutes Some intense 

flashing 0 1/4 inch

Lavatory fluid 16 minutes Some intense 

flashing 0 3/8 inch

Poly-X 

(BMS13-42)

1 percent 

NaCl 5 minutes Some intense 

flashing 0 1/2 inch

1 percent 

NaCl

Less than 

1/2 minute

Some intense 

flashing 0 1/4 inch

1 percent 

NaCl

About 3 

minutes Strong arcing 1 (reset, then 3) 1/2 inch

Kapton 1 percent 

NaCl

About 1.5 

minutes Strong arcing 3 (reset) 2 inches

1 percent 

NaCl

Less than 

1/2 minute Strong arcing 3 (no reset) 7/8 inch

303 Oscillogram measurements taken during one of the pretests indicated that current peaks of 75 amps

occurred and more than 350 J of electrical energy was dissipated. Subsequent examination of the wire

bundle revealed about 1/4 inch of damage and a char buildup around the cut in the insulation; both

conductors were still intact. 
Factual Information 147 Aircraft Accident Report

Lectromec also conducted wet dielectric tests304 on the wires after seven of the wet

short-circuit tests to determine whether the wires that had not been intentionally cut had

been damaged. Multiple wires within the bundle failed six of the seven wet dielectric tests. 

Metal Shaving Abrasion Tests (Dry)

Most of the metal shaving abrasion tests were conducted in two experimental

configurations305ótwo tests were conducted with the wire bundle secured to surfaces such

that the bundle had a 90°-angle bend in it, and four tests were conducted with wires in a

common bundle, aligned longitudinally with each other. In both configurations, one of the

surfaces to which the wires were attached was stable, and the other surface was movable.

The movable surface was attached to an ìoscillator,î which had a 1/2-inch movement

stroke. When activated, the oscillator moved the wires to which it was attached. 

The two 90°-angle metal shaving abrasion tests were conducted with a seven-wire

bundle and a metal shaving at least 0.5 inch long woven between the wires (one bundle

had a steel shaving, the other had an aluminum shaving). The bundles were secured to a

stationary platform at one end and to a movable bar at the other end, in a manner that

resulted in a 2-inch radius, 90° bend in the wire bundle. The linear motion of the oscillator

caused the wires in the bundle to squeeze together with the metal shaving and then relax

repeatedly. These tests were run concurrently, for 19 1/4 hours, with no electrical events

(that is, scintillations, flashing, or arcing) observed and no circuit breakers tripped.

Post-test examination revealed some damage to the wires. Damage observed where the

shavings had been woven into the wire bundles appeared to be mostly superficial damage

to the insulation topcoat (with the steel shaving resulting in slightly more damage).

Lectromec reported that the damage resulting from the flexing and moving of the

25+ year-old wires in the bundles was more severe, including broken and cracked wires,

with exposed conductors in both bundles; these wires subsequently failed wet dielectric

tests. 

During the four longitudinal metal shaving abrasion tests, two identical wire

bundles (four wires in each) were secured together with nylon ties along their midsections,

and several metal shavings were placed between the two bundles. Both ends of one of the

bundles were then secured to a stationary platform, while the ends of the other bundle

were secured to a movable bar. The results of these tests varied from cases in which the

304 Wet dielectric tests provide an indication of the integrity of a wireís insulation. These tests were

conducted by placing individual wires in a water bath with both ends out of the water, then applying

1,000-volt d.c. to the wire conductor while the water bath was grounded. If a current was measured in the

circuit, it was an indication that the wire had short-circuited through its insulation to the water bath, and the

wire failed the test. 

305 Several different configurations of drill shaving abrasion tests were considered and tried before

Lectromec decided to use the two configurations described in this section. Lectromec reported that during

these early trial tests, the shape and size of drill shavings had a significant effect on the test results. For

example, in most cases thin aluminum shavings often broke apart and fell out of the wire bundles and did not

interact with individual wires, whereas larger metal shavings resulted in the tripping of circuit breakers.

Further, thin metal shavings often acted as fuses and evaporated after carrying an initial current between

wires (thus ending the short-circuit condition), whereas shavings with relatively large cross-sections could

support a larger current without evaporating. 
Factual Information 148 Aircraft Accident Report

wire bundle shed the metal shavings without a short circuit, to others in which the metal

shavings cut through the insulation, exposed the conductor, and caused a transfer of

energy to occur. For example, one 2.5-millisecond pulse melted the metal shaving without

tripping circuit breakers, whereas another test resulted in prolonged flashing events with

power peaks of greater than 10 kW, damaging all the wires in the test bundle. Post-test

examination of the wires revealed that bundles containing aluminum alloy shavings were

less damaged by contact with the shavings than those containing steel shavings. However,

as with the 90°-angle metal shaving abrasion tests, movement of these wires resulted in

cracked and damaged wires that subsequently failed wet dielectric tests. Examination

revealed that cracks had appeared in the insulation. No circuit breakers were tripped

during these tests. 

Dry Short-Circuit Tests, with Metal Shavings

Eight tests were conducted to document the dry short-circuit capability of the

metal shavings when a Poly-X wireís conductor was exposed. In these tests, two of seven

wires in a bundle were cut circumferentially (about 1 mm wide) to a depth that exposed

the conductor, placed on opposite sides of the bundle, and positioned so that the exposed

conductor was visible. Aluminum or steel metal shavings of various thicknesses were then

woven through the bundles so that they touched both of the exposed conductors. The

bundles were then suspended horizontally in the test chamber and attached to the

generator, and 120-volt a.c. power was applied. 

According to the Lectromec report, all but one of these tests resulted in short

flashes of various intensity, none of which tripped a circuit breaker. Strong arcing was not

observed in any of the eight tests of Poly-X wire, but one test with a steel shaving resulted

in what Lectromec described as an ìintense flash.î During four of the tests, a visible flash

was observed and then the sample became dormant (no current flow was measured). In

these four cases, minimal or no damage was observed to the wire insulation; all

(non-predamaged) wires passed wet dielectric tests after these tests. However, the other

three dry short-circuit tests resulted in flashing events of longer duration and in additional

damage to adjacent wires after the flash/arc event. After three of the eight tests, several

(one to five) non-predamaged wires failed wet dielectric tests. Oscillogram measurements

indicated that the peak current during these tests was 140 amps and that the peak electrical

energy dissipated was 18 J. 

1.16.6.2.1 Circuit Breaker Protection Information 

According to the USAF Aircraft Mishap Investigation Handbook for Electronic

Hardware, the typical thermal circuit breaker contains a bimetallic strip and two electrical

contacts (one of which is spring-loaded) in a circuit. When the circuit is heated, the

electrical contacts heat the bimetallic strip, causing it to bend. The amount of bending is a

function of the amount of heat, which is dependent on the time and amount of current

flowing through the circuit. When the heat exceeds a preset amount, the bending of the

bimetallic strip causes the release of the spring-loaded contacts to open the circuit. The

circuit can be opened or reset by manually pulling or pushing on the circuit breaker button,

respectively. The position of the button denotes the state of the circuit breaker. Figure 34
Factual Information 149 Aircraft Accident Report

shows a ìtrip-curveî chart for a typical circuit breaker (excerpted from Texas Instruments

KlixonÆ specification), which depicts the relationship between the amount of current that

can pass through a circuit breaker and the amount of time until the circuit breaker opens. 

The USAF handbook also describes typical circuit breaker failure modes,

including welding and erosion/wear of the electrical contacts, material transfer,

overcurrent, and contamination. Further, the handbook states that ìthe mechanical

structures have changed over many years; therefore, age is also a factor.î 

Boeing Document D6-40359 describes a study, ìImpact of Electrical Overloads on

Wire andÖCircuit Breakers Used on 707, 727, and 737 Airplanes,î dated February 2,

1972. The document stated that ìearly in 1971, [an air carrier] reported a failure of

aÖcircuit breaker. As a result of this failure and subsequent investigations, there has been

consideration given to the need for electrically checking all [similar] circuit breakers

which are in service.î The document further stated that circuit breakers that have been in

service ìcould not reasonably be expected to meet the overload current versus trip time

calibration curves of a new breakerî and contained a description of how the bimetallic

strips may lose integrity and degrade with exposure to high internal temperatures. 

Subsequently, Boeing studied about 1,200 circuit breakers that had been removed

from service during a fleetwide survey306 and found that ìinitial results of on-board testing

showed that 72 percent of all breakers tested satisfied the procurement specifications.î

Further laboratory testing of the 28 percent that did not meet procurement specifications

Figure 34. A trip-curve chart for a typical circuit breaker (excerpted from Texas 

Instruments KlixonÆ specification).
Factual Information 150 Aircraft Accident Report

revealed that an additional 9 percent of the circuit breakers originally tested (for a total of

81 percent) satisfied the procurement specifications, despite initial results indicating that

they did not. Boeing speculated that the test equipment and methods used during the initial

tests might have resulted in erroneous failure indications for the 9 percent. Further, Boeing

speculated that the test equipment and methods used in the initial tests might have

damaged circuit breakers within the remaining 19 percent that did not satisfy procurement

specifications. After Boeing revised the pass-fail criteria and retested the circuit breakers,

test results showed that 99.2 percent of the tested circuit breakers could function in time to

protect the wires. 

After a European 757 experienced an arc-tracking event in Kapton wiring on

May 10, 1985, Boeing investigated methods to protect against arc-tracking events. Boeing

document D045Y41102TR, ì270 [volt d.c.] Wet Dielectric Arc Tracking Tests,î described

development of a 270-volt-d.c. trip curve that would provide protection equivalent to that

provided by a thermal circuit breaker with an a.c. power source. Boeing found that starting

loads and in-rush current required circuit protection levels to be set higher than the rating

of available circuit breakers. 

Boeing discovered that the trip characteristics to interrupt arc-tracking events

below the 1,000 percent (rated) current level of d.c. circuits were the same as those of a.c.

circuit breakers.307 Therefore, Boeing designed an electronic circuit breaker to eliminate

the arcing problems associated with d.c. switches. However, Boeing also found that

application of 270-volt d.c. electricity to 115/200-volt a.c. circuit breakers resulted in the

melting of the breaker contacts. 

To offer additional protection to electrical circuits against the unwanted effects of

electrical arcing, nonaerospace manufacturers have begun incorporating arc-fault

detection technology into products known as arc-fault circuit breakers (AFCB, also known

as arc-fault circuit interruptors). According to UL,308 an AFCB is a device intended to

reduce the number of arcing-fault fires by opening the electrical circuit when an arc fault

is detected. An AFCB differs from a typical thermal-based circuit breaker in that the

AFCB has complex electronic circuitry that can identify specific characteristics or

signatures of the current or voltage waveform that are unique to electrical arcing. In 1998,

in conjunction with the National Electrical Manufacturing Association, UL began working

with manufacturers and other interested parties to form an Industry Advisory Group to

develop a standard for safety for AFCB products. In February 1999, UL published the

First Edition of the Standard for Safety for Arc-Fault Circuit Interruptors. The USAF,

FAA, and Boeing have been working to develop AFCBs for aerospace use, but no AFCBs

are currently certificated for use in airplanes. 

306 The 1,200 circuit breakers were removed from 50 airplanes and were considered representative of the

circuit breakers in the fleet.

307 Current may briefly exceed 1,000 percent, as shown in figure 34. 

308 This paragraph is reprinted with the permission of UL, from ìAFC[B]s Show Promise to Save Lives

by Preventing Electrical Fires.î 1999. On the Mark, Vol. 5. No. 3-4.
Factual Information 151 Aircraft Accident Report

1.16.6.3 Aging Effects on Material Properties of Wires 

To better understand the condition of wiring in aging transport-category airplanes,

the Safety Board contracted with Raytheon Systems Company (Raytheon) to test aged

wire samples to the specifications for newly manufactured wire of that type. The aged

wire samples were similar in age and type to the wiring on TWA flight 800 and were

obtained from three recently (not more than 6 months before the wire samples were

obtained) retired airplanesótwo 747s, one manufactured in 1970 and the other in 1973,

and a Douglas DC-10 manufactured in 1973.309 The wire samples ranged in length from

1 foot to more than 26 feet and in size from small, partial bundles with just a few wires to

large bundles with many wires. The samples were taken from pressurized fuselage

locations, most of which were relatively protected from traffic, exposure to light, or other

potential environmental stresses. Figures 35a, 35b, and 35c show the sample origination

locations for the 747 manufactured in 1970, the DC-10, and the 747 manufactured in

1973, respectively. Table 9 further describes the origin and potential environmental

exposure of each sample.

Table 9. The origin and potential environmental exposures of test samples.

309 According to the Raytheon report, the wires tested were ìessentially identical.î Raytheon described

all of the tested wires as tin-coated copper conductors with a dual layer of cross-linked extruded

alkane-imide polymer insulation and a topcoat or Poly-X type wire. There was only one manufacturer

(Raychem) qualified to produce wire for either the Boeing or MIL specifications during the wire usage

period. When Raychem stopped producing this wire type in 1977, the associated MIL specification was

cancelled. 

Aircraft Sample Origin of Wire

Potential 

Environmental 

Exposure

747, Delivered 1973 Samples 1, 2, 3, and 4 Flight engineerís panel in 

cockpit

Possible ultraviolet 

(UV) ray exposure

Sample 5 STA 380, right side of flight crew 

exit door

Possible UV ray 

exposure

Sample 6 STA 1438, directly beneath floor 

panel of seats 37A, B, and C Benign

Sample 7 Just forward of the CWT on the 

right side Benign

DC-10, Delivered 1973 Samples 1 and 3 Cockpit overhead center panel Hot, dry, no UV ray 

exposure

Sample 2 STA 770, a 12-foot long section 

from above the R2 door Benign

Sample 6 Not marked Not known

Sample 7 Underneath flooring in cockpit Benign

Sample 8 STA 919 to 1570, on the right 

side, overhead R2 door Elevated temperatures

747, Delivered 1971 Sample 1 Forward cargo compartment Benign
Factual Information 152 Aircraft Accident Report

STA

90

STA

140

OVERHEAD HATCH

TOP OF

PILOTS 

FLOOR

STA

520

STA

700 STA

830

KEEL

BEAM

ACCESS

DOOR

NO. 1

CARGO

DOOR

STA

134.75

NOSE

CARGO

DOOR

STA

1480

STA

1000 STA

1140 STA

1241

STA

1350

Sample 6

Sample 1

SEC 41 SEC 42 SEC 44 FRONT SPAR MID SPAR REAR SPAR 

Figure 35a. Sample origination locations for the 747 manufactured in 1970.

Figure 35b. Sample origination locations for the DC-10 manufactured in 1973.
Factual Information 153 Aircraft Accident Report

In addition to visual examinations, Raytheon tested the wire samples against the

following performance requirements for new wire: 

ï Mechanical and electrical tests, including concentricity tests (to measure the

thickness of the insulation layers and the overall sample geometry), dielectric

tests (to determine the integrity of wire insulation), insulation resistance tests

(to measure the electrical resistance that the insulation imparts between the

conductor and ground), insulation tensile strength and elongation tests (to

identify materials that have weakened because of thermal breakdown, UV

degradation, or other mechanisms), and notch sensitivity tests (to measure the

ability of a wire insulation to resist the propagation of a nick or cut through the

insulation layers to the conductor). 

ï Thermal tests, including blocking tests (to ensure that the insulation does not

adhere to itself after heating), low-temperature (cold-bend) tests (to ensure that

the wire retains physical and electrical integrity through mechanical and

electrical stress at the temperature extremes), shrinkage tests (to indicate

whether the wire insulation is physically and dimensionally stable when

exposed to short-term high temperature), thermal shock tests (to measure an

insulationís ability to resist shrinkage and expansion following several thermal

cycles to temperature extremes), and wrapback tests (to evaluate the wireís

ability to withstand thermal stress while under mechanical stress). 

STA

90

STA

140

OVERHEAD HATCH

TOP OF

PILOTS 

FLOOR

STA

520

STA

700 STA

830

KEEL

BEAM

ACCESS

DOOR

NO. 1

CARGO

DOOR

STA

134.75

NOSE

CARGO

DOOR

STA

1480

STA

1000 STA

1140 STA

1241

Samples 1, 2, 3, 4

Sample 5

Sample 7

SEC 41 SEC 42 SEC 44

STA

1350

FRONT SPAR 

MID SPAR 

REAR SPAR 

Figure 35c. Sample origination locations for the 747 manufactured in 1973.
Factual Information 154 Aircraft Accident Report

ï Thermal aging tests, including accelerated aging tests (to evaluate a wireís

ability to withstand a higher temperature under mechanical stress for a short

period of time), lifecycle tests (to evaluate a wireís ability to withstand slightly

elevated temperatures under mechanical stress for a longer exposure time), and

electrical integrity tests (to evaluate a wireís electrical integrity during forced

hydrolysis and mechanical properties and electrical integrity during heat

aging). 

According to the Raytheon report, in general, the wire samples obtained from the

three airplanes appeared to be in similar, ìgenerally decentî condition; however, the

condition and performance of the specific wire samples varied within and across the

airplanes. Raytheonís report indicated that visual examination of the wire samples

revealed a variety of conditions: the presence of debris, such as lint and small chips, paint

residue, and possibly foam or adhesive residue; contamination by what appeared to be oil

or grease and a dark brown film; mechanical damage, such as insulation cracking,

indentations caused by tight clamps or nylon tie wraps, abrasion of the insulation, and

various nicks and cuts; and deep hot-stamp marking, which resulted in cracking of the

insulation. The wire insulation was also observed to be brittle in all cases, and, in some

cases, the outer layers cracked as a result of routine handling in the laboratory. About

20 percent of the wires received with damage failed in dielectric testing, indicating that

the wiresí insulation did not meet the integrity requirements for new wire and that fluid

could reach core conductors. 

Further, although the wire samples met most of the performance requirements for

new wire when subjected to tests for insulation resistance, notch sensitivity, blocking,

cold-bend shrinkage, thermal shock, and accelerated aging, problems/failures were noted

during concentricity, elongation, dielectric, and lifecycle tests. For example, the Raytheon

report indicated that several wire samples from different airplanes failed concentricity

tests, ìindicating that the insulation was degrading with layer separation.î Additionally, in

several cases the elongation of the outer insulation was extremely low, indicating that ìthe

wire became more brittle with age....None of the wire samples tested passed the lifecycle

test, indicating that the long-term life of all samples may be limited. Additional thermal

life testingÖshowed that the samples had some limited life remaining before beginning to

fail physically and electrically.î 

The Raytheon report included the following conclusions: 

1. Most wire samples tested had definitely aged as indicated by the outer jacket

[insulation layer] failures in accelerated aging and lifecycle. 

2. The inner insulation continued to provide electrical integrity, as it was

designed to do, although the outer jacket [insulation layer] was losing its

ability to mechanically protect the wire. [See figure 10.] 

3. The wire submitted for testing would be expected to perform adequately in the

short term, provided there was no added stress beyond what it had experienced. 
Factual Information 155 Aircraft Accident Report

4. The weak point in this wire type appears to be the lack of ability to maintain

physical and electrical integrity during extended elevated thermal exposure

with mechanical stress. As the wire ages, there is an increased risk of dielectric

problems [breakdown of wire insulation integrity], mechanical damage and

loss of electrical integrity. 

5. No wire was submitted that was exposed to high levels of environmental

exposure (such as wires from wheel wells, wing leading edges, etc.). Wire

from these areas, would be expected to perform worse than the wire submitted.

[See figures 35a, 35b, and 35c for sample locations.]

Investigators found original Boeing markings (such as green W42A lettering) on

the recovered wiring for most of the accident airplaneís systems,310 and many of the

recovered FQIS wiring throughout the airplane had the darkened appearance of similar

wire materials found in similar-aged airplanes. (As previously stated, the accident airplane

had 93,303 total hours of operation at the time of the accident.) 

Raytheon found wires other than Kapton and single- and dual-layer insulation

Poly-X in the sample wire bundles that it examined. These wires were insulated with

single- and dual-layer Teflon and cross-linked polyalkene/polyvinylidene fluoride (an

earlier type of wire that had been manufactured per MIL-W-81044/9).311 The Safety Board

extended its contract with Raytheon to include testing of these additional three wire types

and some various connectors found with the wire samples.312

As with the previous wire tests, Raytheon tested these materials to the acceptance

standards for new material. Raytheon indicated that although the wire samples were in

ìgenerally decent condition for having 25 years service life, there were problems evident

during the inspection and performance testing.î313 Although these wire samples passed

most of the tests to which they were subjected, test failures were found. For example, the

wire sample with a single layer of Teflon insulation exhibited two dielectric failures, the

wire sample with a dual layer of Teflon of insulation failed the low-temperature

(cold-bend) tests, and the MIL-W-81044/9 wire sample failed the low-temperature

(cold-bend) and lifecycle tests. Specifically, Raytheonís report regarding the second set of

wire tests concluded the following: 

ï The three wire types evaluated have degraded in different manners during

service life. 

310 TWA personnel installed some galley, lighting, and navigation/communication avionics wiring

shortly after the airline took possession of the airplane and again during a modification in 1992. 

311 A few short lengths of another wire (BMS13-38) were also found, but insufficient wire was available

for testing. According to a Raychem technical paper presented at the 18th International Wire and Cable

Symposium in Atlantic City, New Jersey, from December 3 to 5, 1969, Poly-X wire provided improved

mechanical strength and fluid resistance at elevated temperatures, whereas ìsome of the newer hydraulic

fluids attacked the MIL-W-81044 wires at elevated temperatures.î 

312 For additional information, see Systems Group Chairmanís Factual ReportóAddendum for Aging

Aircraft Wire Testing by Raytheon, Report 2, dated January 28, 2000. 

313 Although Raytheonís statement indicated that the wire samples had 25 years of service life, some of

the wires tested had not been available for use for that long. 
Factual Information 156 Aircraft Accident Report

ï Single layer BMS13-48 [XL-ETFE, such as Tefzel] type wire developed

dielectric failures, possibly due to the thinner wall thickness providing less

mechanical protection to the wire. 

ï Dual layer BMS13-48 [XL-ETFE, such as Tefzel] type wire passed dielectric

tests without failures, but the outer layer has lost strength and flexibility as

evidenced by the results of tensile, elongation, and low temperature cold bend

tests. 

ï Dual layer MIL-W-81044/9 type wire failed two tests. The outer layer cracked

during the low-temperature cold bend test even though retaining electrical

integrity. The wire failed lifecycle tests, indicating that the wire has aged

significantly and has limited additional life.

ï The BMS13-48 [XL-ETFE, such as Tefzel] wire appeared to be in better

physical shape than the [BMS13-42 Poly-X] wire tested [previously].

ï None of the wire samples tested had been exposed to high levels of

environmental stresses. Wire from areas exposed to high levels of

environmental stress (such as wires in landing gear wheel wells, wing leading

edges, etc.) would be expected to perform worse than the wire submitted.

ï The connectors from all three aircraft showed signs of aging, such as the

presence of corrosion and hardening grommets.

ï Connectors from all three aircraft passed dielectric testing indicating they

insulated the contacts from each other and the shell, but several of the

connectors did not have the ability to have maintenance performed on the

connectors. Contacts were not retained, grommets began to crumble, and some

contacts were not removable. 

1.16.6.3.1 Previous Reports about Wire Degradation and Aging

In May 1976, Boeing published document D6-42974, ìInvestigation of Wire on

Wire Abrasion Testing,î in which it documented the wear characteristics for several wire

types. The results indicated that the resistance of the tested wire types was rated as

follows, from best to worst: BMS 13-51 (Kapton), BMS13-13 (PVC Nylon), BMS13-42C

(Stilan), BMS13-31 (mineral-filled PTFE), BMS13-48 (Tefzel), and BMS13-42A/B

(Poly-X). Boeingís report stated that ìthe results are more or less as would be expected

and agree with predictions based on in-service experience by Boeing and others and with

other evaluations of these wires.î 

Reports from NAFI noted that Poly-X was adopted for use on January 2, 1970,

because MIL-W-81044/16 and /18 were phased out. Poly-X was first installed in the F-4N

Phantom II, and the first wire insulation malfunctions were recorded in

September 1975.314 The NAFI report also stated that failures of Poly-X in F-14 aircraft

had been reported about 3 years after installation. On the basis of its observations of

Poly-X failures, NAFI established that the time to field failures for Poly-X in U.S. Navy

314 For additional information, see U.S. Naval Avionics Facility. 1977. Investigation of Poly-X Wire

Deterioration on F-4N Aircraft. NAFI Report No. 01-76, NARF-NORIS. 
Factual Information 157 Aircraft Accident Report

operations was between 3 and 5 years in service.315 Additional laboratory tests conducted

by NAFI and historical evidence indicated that Poly-X wires aged 2.3 to 7 times faster

than Kapton. The NAFI report stated, ìif it is assumed that there is reasonable correlation

between laboratory tests and service experience, then Kapton-insulated wire may fail after

7-21 years of service. The use of milder cleaning compoundsÖshould extend the service

life of the installation.î 

After the NAFI reports were published and the U.S. Navy switched to milder

cleaning compounds, the military, the aviation industry, and Boeing continued to study the

wire-aging process. A 1997 Lectromec report316 documented the aging of Kapton wire

insulation in various locations in U.S. Navy P-3 airplanes. This report indicated that wire

insulation in areas exposed to sunlight and moisture could reach the end of its service life

(chemically) within 1 year, whereas wire insulation in protected areas in the same airplane

could survive for up to 10 years. Figure 36 shows the percent probability of failure of

Kapton wiring in different locations in the P-3 airplane. 

The Safety Board also found that Boeing documented and described cracked wire

insulation in SL 747-SL-20-048, issued January 25, 1995. Boeing described ìthe virtue of

a wire by wire inspectionî of wing leading edge and engine strut areas. However, the SL

also stated that ìBoeing generally does not recommend special wiring inspections unless a

fault is experienced or an area is disturbed for other reasons.î

315 The Safety Board notes that airplanes flown by the U.S. Navy may experience environmental

conditions to which civilian airplanes may not be subjected, including continuous operations in humid

salt-air environments and frequent washings with acidic cleaning solutions. Further, wires in the wing-fold

area (which are nonexistent on civilian airplanes) were often exposed to more vibration and environmental

conditions (such as sunlight, wind, rain) than wiring in a like location on a civilian airplane. 

316 For additional information, see the following Lectromec reports/papers: Aging Measurements of

Operating Aircraft Wiring; Implications for Specification Writing and Aircraft Reliability, dated July 10,

1997; and Description of Lectromec WIDAS Aging Support Program, dated November 11, 1998. Also, see

Eaton, D. R. and Bruning, A. 1999. Economic Aircraft Wire Maintenance, Inspection and Repair:

Performance Implications. 
Factual Information 158 Aircraft Accident Report

Figure 36. Chart showing the percent probability of failure of Kapton wiring in different 

locations in the P-3 airplane.

Location

Bomb Bay

Wing, Outboard Trailing Edge

Galley/Aft Cabin

Wing Center Leading Edge

Forward Electrical Load Center

Avionics Bay C1

Wing Inboard/Ro Leading Edge

Avionics Bay H1

Hydraulic Service Center Under De

Main Wheel Well

Nose Wheel Well

Wing Center Trailing Edge

1 2 5 10 20

0

0

0

0

0

0

15

21

20

38

31

0

0

0

0

0

0

0

20

23

26

42

57

74

0

0

0

15

24

43

32

40

39

50

89

91

24

28

41

23

35

57

46

46

56

72

100

100

33

53

61

30

48

68

60

78

64

100

100

100

Years

Percentage of Failures
Factual Information 159 Aircraft Accident Report

1.16.6.4 Electromagnetic Interference 

1.16.6.4.1 Electromagnetic Interference from External Sources 

The Safety Board requested that the DoD Joint Spectrum Center (JSC) conduct

research into and calculate the electromagnetic environment (EME) in which TWA flight

800 was likely to have been operating at the time of the accident (that is, the

electromagnetic field strength from the external emitters to which the airplane was

exposed during its flight). Fixed and mobile radio frequency (RF) emitters317 in the

TWA flight 800 environment were identified. Ground-based emitters with frequency

assignments318 for the eastern United States (as identified in the JSC-maintained

frequency assignment databases) comprised the fixed emitters. These ground-based

emitters included the following equipment types: ASR, air route surveillance radar, airport

surface detection equipment, military navigation radar, weather radar, terminal Doppler

weather radar, long-range tracking radar, U.S. Navy radar shore installations, space object

tracking radar, over-the-horizon radar, and research and development emitters.319 The RF

equipment on board air and sea platforms near TWA flight 800 (as identified by the Safety

Board, FAA, U.S. Navy, USCG, and Air National Guard records) comprised the mobile

emitters. 

After the RF characteristics of the fixed and mobile emitters were collected, the

electrical field strengths and energy densities these emitters could generate at the accident

airplaneís position and altitude were calculated for the time of its last secondary radar

return. The calculations were based in part on the emitterís distance from that location and

the propagation path losses associated with those distances. The JSC report indicated that

the following assumptions were used in calculating the potential electric field strengths at

the TWA flight 800 location: 

All fixed and mobile transmitters were assumed to be emitting except commercial

airline high frequency (HF) communicationÖequipment. Commercial airline HF

[communication] equipment are typically not employed whenÖVHF air traffic

control is available.[320]

 Military HF [communication] transmitters were assumed

to be active at the time of the accident.

In calculating the peak and average electric field strengths generated at TWA

flight 800, maximum values of peak power, duty cycle, and antenna mainbeam

gain were assumed. Where the duty cycles of pulsed emitters were unavailable,

317 According to an April 13, 2000, letter from the JSC, emitters are ìpulsed sources of low duty cycle

and narrow antenna beamwidths.î 

318 The agencies responsible for ensuring that there are no conflicts between different emitters (the

Federal Communications Commission for civilian emitters, the JSC for military emitters) assign frequencies

that can be used for various functions. 

319 Although television, radio, and cellular telephone ground emitters also contribute to the EME, the

electromagnetic field strengths and energy densities generated by these emitters are significantly weaker

than the other emitters considered during this research. 

320 Because commercial airline HF equipment is typically not used and to limit the scope of the JSC

analysis, it was assumed that this equipment was not emitting at the time of the accident. 
Factual Information 160 Aircraft Accident Report

the JSC compared combinations of pulsewidths and pulse repetition rates to

determine the highest duty cycle. 

The avionics on board USAir[ways] flight 217 [located about 2.5 nm southwest of

TWA flight 800 at 17,000 feet msl] were assumed to be representative of the

commercial aircraft in the vicinity of TWA flight 800. Since USAir[ways] flight

217 was the closest commercial aircraft to TWA flight 800, the electric field

strength predictions due to the emitters on board USAir[ways] flight 217 were

assumed to represent the worst case for the EME due to commercial aircraft

emitters. 

[To reduce the size of the list of emitters to be further analyzed,] a peak electric

field strength of 1 volt/meter (V/m) at the accident location was selected as a

conservative threshold for the minimum field strength that needed to be

considered.[321]

The JSC report stated that when it included terrain and antenna pattern data in its

electric field strength calculations, it identified 40 fixed ground-based emitters that had the

potential to generate peak electric field strengths greater than 1 V/m at the accident

location. Two of the 40 fixed ground-based emitters were identified as having the potential

to generate an average electrical field strength322 greater than 1 V/m at the accident

location: the Westford, Massachusetts, tracking radar and the Brookhaven, New York,

weather radar emitters. According to the JSC report, its calculations revealed that no

mobile emitters, shipboard or airborne, could generate average electric field strengths

greater than 1 V/m. The JSC report further stated, ìalthough many classified emitters were

identified in the environment, none of them were close enough or powerful enough to

generate peak or average electric field strengths greater than 1 V/m.î 

The energy density values (in mJ/m2) were calculated to help relate the field

strength values to units which the investigators could more readily compare with the MIE.

According to the April 13, 2000, letter from the JSC, it is possible for the energy density

from emitters to overlap in time and space, having an additive effect on the energy density

available at that time and space. However, according to JSC, the probability of the energy

density from any two emitters to precisely overlap in time and space at the accident site

was less than 1.85 x 10-4 and would result in an energy density increase (for the highest

energy density emitter) from 0.013 mJ/m2 to 0.02 mJ/m2. Further, the likelihood of

additional emitters overlapping in time and space at the accident site decreased

exponentially with each additional emitter. For example, the calculated probability of

3 dominant emitters overlapping in time and space at the accident site was 7.1 x 10-9 (with

a resultant energy density increase to 0.024 mJ/m2

), whereas the probability of 11

dominant emitters overlapping in time and space at the accident site was less than

1.1 x 10-83 (with a resultant energy density increase to 0.043 mJ/m2

).

321 The 1 V/m threshold was used based on the judgement of JSC engineers that there was no way to

induce an explosion in the CWT with field strengths under that level. Subsequent NASA Langley Research

Center (LaRC) research confirmed this engineering judgement (see section 1.16.6.4.1.1). 

322 Average field strengths are the ones that would better create the energy required to create a spark. The

peak field strengths exist for such a short period of time (microseconds) that there is not enough time to

affect wires in a way that would create a spark with enough energy to ignite a fuel/air mixture. 
Factual Information 161 Aircraft Accident Report

1.16.6.4.1.1 NASAís Report Regarding Electromagnetic Interference from 

External Sources

At the Safety Boardís request, NASAís LaRC conducted tests (using the data

developed by the JSC) to determine the worst-case energy levels from external RF

transmitters that could be induced to fuel tank FQIS components. To this end, LaRC

researchers developed and validated a new computer code that was used to predict the

passenger cabin shielding effectiveness and to model signal-to-wire coupling within the

cabin.323 The researchers used a dominant external emitter identified by JSC324 as the input

signal for their computer code. The LaRC report indicates the following:

the maximum available energy inside the TWA 800 passenger cabin from any

identified dominant emitter was determined to be less than 0.1 [mJ], which is

below estimates of the minimum ignition energy required for ignition of Jet A fuel

vapor. Energy coupled to aircraft wiring within the passenger cabin will be several

orders of magnitude (approximately 10-4 times) less than this. 

1.16.6.4.2 Electromagnetic Interference from Internal Sources 

The Safety Board contracted with the NAWC-AD to measure the voltage

harmonics on the captainís channel of the CVR325 and determine if electrical power

transients could induce sufficient energy into the FQIS wiring through capacitive and

inductive cable coupling to ignite fuel in a 747-100 airplaneís CWT.326 In November 1999,

NAWC-AD conducted electrical ground tests at AAR Aircraft Services, Inc.ís, facilities in

Roswell, New Mexico. 

The tests in Roswell, New Mexico, were inspired by laboratory tests conducted at

Boeing in 1997,327 which created transient voltages in CWT FQIS wires by switching on

and off the power (simulating large power relays or motor loads) in wires that had been

laid parallel to the CWT bundle. The tests induced up to 0.6 mJ of energy into a CWT

harness, exceeding Boeingís specification of 0.02 mJ and the MIE of 0.25 mJ referenced

323 Note: The EMI would be induced on wires outside the CWT, and the energy would then be

transferred into the CWT. 

324 The JSC identified the dominant external emitter as the Brookhaven, New York, weather radar and an

airborne radar emitter as a close second. The emitter used by the LaRC researchers was the search radar at

Riverhead, New York, which had the third highest field intensity levels. This emitter was used to allow for a

decreased computation time for the numerical computer code developed by LaRC. Even at this lower

frequency, the calculations required 182 central processing unit hours to complete. To perform the

calculations using the dominant emitter would have taken on the order of 14,000 central processing unit

hours. 

325 For more information, see section 1.11.1.1.

326 For the purposes of this investigation, the Safety Board examined the effects of EMI as a potential

ignition source for the CWT overpressure event. The Board did not examine the potential effects that EMI

might have on an airplaneís navigation/communication systems. 

327 These tests were described by a Boeing engineer during his testimony at the Safety Boardís public

hearing in December 1997. They are also discussed in the FAAís NPRM 97-NM-272-AD and in Boeingís

May 26, 1998, comments on that NPRM. 
Factual Information 162 Aircraft Accident Report

in API 2003. The tests were conducted using maximum source voltages of 800 volts with

resultant voltage transients in excess of 1,000 volts.328

The Safety Board contracted with NAWC-AD to conduct tests similar to those

conducted by Boeing, except that they were to be conducted on an airplane using actual

airplane systems to generate the switching transients. To prepare for these tests,

NAWC-AD and Safety Board personnel examined and analyzed Boeing drawings and

documents and examined more than 900 wires from systems in a 747-100 to determine the

electrical loads (components) to which they were attached and to determine which of these

loads would have the highest potential for coupling energy to CWT FQIS wiring if their

wires were corouted. The wires connected to loads with the highest coupling potential

were repositioned within wire bundles/trays so that they were adjacent to the FQIS

wiring329 to provide a ìworst-caseî routing situation. 

Because the CWT access doors in the 747-100 did not provide adequate access for

instrumentation for these tests, the researchers used a surrogate CWT330 that provided

better instrumentation access. The surrogate CWT was positioned on the ground by the

test airplaneís main landing gear, and the airplaneís FQIS wiring was disconnected from

the airplaneís CWT and connected to the surrogate CWT. The airplaneís systems were

powered using engine-driven generators, and the FQIS was energized.331 Voltage and

current measurements were taken at fuel probes in the surrogate CWT as the airplane

systemsí electrical loads were cycled. Additional measurements were taken with

conductive debris (a few strands of aluminum wool) placed either across terminals of a

fuel probe or from a terminal of a fuel probe to the simulated CWT to create and calculate

the energy dissipation through the debris. According to the NAWC-AD report,332 the

maximum energy calculated for a transient through conductive debris (a few strands of

aluminum wool) was 0.125 mJ. 

328 The 800 volts were much higher than the 25 to 125 volts found in wiring corouted with the 747 FQIS

wiring system; however, it prompted further tests. 

329 The wires for the high coupling potential loads were only moved to be adjacent to the FQIS wiring

when they were corouted in the same bundle or tray. In portions of their runs where the two sets of wires

were not corouted, they were not moved. 

330 According to a NASA report, although the surrogate tank had a similar overall volume to the 747-100

CWT, the surrogate tank had different overall dimensions, was made of steel (not aluminum), and was not

divided into bays like the CWT.

331 To avoid igniting fuel in the airplaneís wing tanks during these tests, nitrogen gas was continuously

pumped into the on-board fuel tanks through the pressure-refueling receptacle on the left wing. The

inertness of the on-board tanks was continuously monitored throughout the tests by measuring the oxygen

content from the outflow vents at the tip of each wing. 

332 See Naval Air Warfare CenterñAircraft Division. 2000. Boeing 747-100 Fuel Quantity Indication

System (FQIS) Susceptibility to Induced Energy from Capacitive and Inductive Cable Coupling.

NAWCADPAX/TR-2000/33. 
Factual Information 163 Aircraft Accident Report

1.16.6.4.2.1 Electromagnetic Interference from Personal Electronic Devices 

On Board TWA Flight 800

The Safety Board was aware of cases in which signals from PEDs reportedly

adversely affected airplane systems.333 Therefore, to determine whether EMI from PEDs

on board the accident airplane may have been involved in this accident, it requested that

scientists at NASA LaRC evaluate the effects of signals from PEDs on board the airplane

(such as laptop computers, video games, AM/FM and other handheld [CB, police, fire,

business, maritime] radios, tape players, compact disc players, cellular telephones, garage

door openers, remote control toys, and pagers) on airplane wiring. The report noted the

following:

[the electromagnetic] environment [created by a transmitter] is highly dependent

upon distance from the transmitter. This relationship changes as the inverse of the

distance squared (1/R2

). For example, if you move from 1 meter to 100 meters

away, the available power [from a transmitter] is reduced by 10,000.

Alternatively, it can be demonstrated that a portable radio transmitting about 5

watts may generate field levels in excess of 100 Volts/meter very close to its

antennaÖ[PEDs] are typically operated over a wide range of frequencies and

power levels, and can often be set to radiate continuously. When used inside an

aircraft, PEDs can be placed very close to aircraft wiring, and their emissions are

not subject to airframe attenuation as [external sources are]. 

The NASA LaRC research into the effect of PEDs on the FQIS system was

designed to study whether a typical portable device could possibly create an ionization

event (spark) or excessive heating, if all available power is applied directly to the FQIS

wiring. To support this effort, LaRC researchers performed the following analysis and

experimental testing: 

ï determined the worst-case electromagnetic threat that could have been created

by commercially available PEDs in terms of power and frequency; 

ï determined the electrical characteristics of the CWT and FQIS wiring in a

retired 747-100 airplane;

ï removed the fuel probes and FQIS wiring from a retired 747-100 airplane,

installed them in a laboratory test chamber, and conducted tests to determine

the RF power required to achieve a spark or excessive heating of the laboratory

installation; 

ï determined scaling factors required to compare the laboratory with the airplane

FQIS installations; and 

ï assessed existing guidelines and references for understanding how continuous

RF signals may impart sufficient energy to ignite fuel vapors. 

333 For example, a June 29, 1998, NASA Aviation Safety Reporting System report (98-30/10-2) cited

anomalous displays on navigation instruments at both pilot positions of a DC-9 during the approach to

landing. The report indicated that after landing, a flight attendant reported that a passenger ìreceived a pager

call while the airplane was on base leg, and refused to turn it off.î 
Factual Information 164 Aircraft Accident Report

The NASA LaRC report concluded that the strongest electrical signal that could be

coupled to the FQIS from PEDs in the passenger cabin and then cause a spark in the CWT

was about 1.5 watts (about one-tenth of the minimum power level required to induce a

spark from the FQIS).334 This was obtained during laboratory testing by applying a

simulated intermittent fault (an arc-gap between the FQIS components and a ground

lead)335 at a worst-case location for voltage enhancements. Without an intermittent fault,

no detectable discharge event was observed, even when applying more than 10 watts of

RF power. The report further stated that the worst-case coupling from PEDs in the

passenger cabin would be coupled to the FQIS wires, not directly to the tank volume.

Although the FQIS coupled power levels predicted to result from a PED signal vary as a

function of emitter frequency, LaRC researchers reported that the maximum predicted

power level was about 0.06 watt. 

1.16.6.5 Boeingís Center Wing Fuel Tank Ignition Fault Tree Analysis336

Computations 

Although the FAA regulations regarding the certification of new airplanes in

December 1969, when the 747-100 airplanes were certificated, did not require

manufacturers to conduct a quantitative evaluation of probability levels for system

failures, the FAA currently does require manufacturers to perform such an evaluation for

certification. The regulatory guidelines addressing quantitative evaluation of failure

modes were adopted in April 1970 as amendment 25-23 to 14 CFR Part 25; the

quantitative evaluation procedures were subsequently detailed in AC 25.1309-1A. The

probability levels are described in terms of the consequences of minor, major, and

catastrophic failures. For example, Section 25.1309 states the following: 

(b) The airplane systems and associated components, considered separately and in

relation to other systems, must be designed so tható 

(1) the occurrence of any failure condition which would prevent the continued

safe flight and landing of the airplane is extremely improbableÖ.[337]

334 For these tests, the NASA LaRC researchers were limited by their laboratory equipment to using

continuous wave signals instead of pulsed signals. Power measurements were used to characterize

continuous wave signals, and they were proportional to energy measurements over the short term. Energy

measurements from continuous wave signals were a function of how long the signal was maintained and

would not have represented the conditions present during the accident. 

335 A ground connects electrical circuits to a large common conductor considered to be at zero electrical

potential (such as the earth or, in the case of an airplane, the skin of the airplane and then overboard to the

atmosphere).

336 A fault tree analysis is designed to examine an end event through consideration of assumed

precipitating events. Each of these precipitating events is in turn broken down until a level is attained in

which no additional precipitating events will occur. The total of all of the events and the way in which they

are tied together creates the fault tree analysis. 

337 AC 25.1309-1A defines the term ìextremely improbableî failure conditions as conditions that are ìso

unlikely that they are not anticipated to occur during the entire operational life of all airplanes of one type.î

These type of conditions are further defined as ìthose having a probability on the order of 1 x 10-9 or lessî

for each flight hour based on a flight of mean duration for the airplane type.
Factual Information 165 Aircraft Accident Report

(d) Compliance with the requirements of paragraph (b) of this section must be

shown by analysis, and where necessary, by appropriate ground, flight, or

simulator tests. The analysis must considerñ

(1) Possible modes of failure, including malfunctions and damage from external

sources;

(2) The probability of multiple failures and undetected failures;

(3) The resulting effects on the airplane and occupants, considering the stage of

flight and operating conditions, and

(4) The crew warning cues, corrective action required, and the capability of

detecting faults.

Similarly, Section 25.901(c) requires the following:

[f]or each powerplant and auxiliary power unit installation [including the fuel tank

system], it must be established that no single failure or malfunction or probable

combination of failures will jeopardize the safe operation of the airplane except

that the failure of structural elements need not be considered if the probability of

such failure is extremely remote. 

Although no fault tree analysis report was required (nor was one conducted) for

the accident airplaneís certification, as part of its involvement in the Safety Boardís

investigation of the TWA flight 800 accident, Boeing developed and submitted a CWT

ignition fault tree ananlysis report on November 25, 1996; this report was subsequently

revised on December 20, 1996.338 The assumptions made in developing Boeingís fault tree

analysis report included the following:

ï Only events considered germane to the TWA event were quantified.

ï Aircraft was assumed to be in climb/cruise configuration.

ï The CWT FQIS worked normally on the previous flight.

ï The delivery schedule of this report required that many of the failure rates

developed for events in this report be subjectively established. Using the

qualitative criteria contained inÖAC 1309-1A as shown below, best

engineering judgement used qualitative risk assessment and then assigned [the

following] quantitative values: probable = 1 x 10-05; improbable = 1 x 10-06 to

1 x 10-09; and extremely improbable = 1 x 10-10. 

ï Quantitative rates assigned were the result of experience with [the] system and

discussion with subject matter specialists.

ï Failure rates will be reviewed for refinement as the final fault tree is

constructed.[339]

ï Air conditioning packs wereÖintact at impact.

338 For additional information, see Systems Group Chairmanís Factual ReportóAddendum for Fault

Tree Data, dated May 19, 2000. 

339 After the December 1996 revision, no further fault tree analysis reports were received from Boeing. 
Factual Information 166 Aircraft Accident Report

ï A combustible fuel/air mixture existed in all fuel tank air spaces as modeled.

ï Scavenge pump switch was turned off at the time of the event. 

Boeingís CWT ignition fault tree analysis report identified the ignition of the CWT

as the top-level failure mode, and the impact of all lower-level elements that compose the

tree was evaluated using mathematical probability relationships. The fault tree analysis

was constructed downward from the top-level event by progressively determining the

conditions required for each lower-level event. Eventually, the tree reached a level where

an event had no preceding events; this event was termed a ìbasic element.î Boeingís CWT

fault tree analysis report identified 167 basic elements. According to Boeing, each basic

element was assigned a failure rate and an exposure time (the maximum time presumed by

Boeing that the condition could go undetected) based on either in-service data or

engineering estimates. These failure rates and exposure times were combined to produce a

probability of failure for each element, using the following formula: probability of failure

= 1 x e(-failure rate x exposure time). Some of the basic elements identified in Boeingís fault tree

analysis report are shown (with their failure rates, exposure times, and calculated event

probability) in table 10.

Table 10. Several fault tree analysis basic elements, as identified in Boeingís fault tree 

analysis report, with their failure rates, exposure times, and calculated event 

probability.

Basic Element

Failure Rate 

(events/hour)

Exposure 

Time 

(hours)

Basic 

Element 

Probability

Fuel/air mixture in CWT will support explosion N/A N/A 1 x 10-00

Metallic object capable of creating spark in CWT 1 x 10-06 8.0 8 x 10-06

Tank sealant decays, resulting in exposed metal surface 1 x 10-07 8.0 8 x 10-07

Electrical faults in surge tank 1 x 10-07 8.0 8 x 10-07

Fire propagates from surge tank to CWT 1 x 10-06 8.0 8 x 10-06

Conductive material bridging terminals 3.55 x 10-07 8.0 2.64 x 10-06

Unique wire-to-wire faults route to FQIS leads in right wing 2.44 x 10-07 8.0 1.95 x 10-06

Fuel probe contacts CWT structure 3.55 x 10-07 8.0 2.84 x 10-06

Unique wire-to-wire faults route to FQIS leads in left wing 2.44 x 10-07 8.0 1.95 x 10-06

Fire propagates from WCS dry bay to CWT 1 x 10-05 8.0 8 x 10-05

Combustible fuel/air mixture in air conditioning pack bay N/A N/A 1 x 10-00

Explosion-proofed equipment does not operate 

as designed

1 x 10-08 60000 6 x 10-04

Fire propagates from air conditioning pack bay to WCS dry 

bay

1 x 10-05 8.0 8 x 10-05

Fuel leak to air conditioning pack bay 1 x 10-05 8.0 8 x 10-05

VSO unit internal fault results in power to FQIS lead 1.4 x 10-07 8.0 1.12 x 10-06

FQIS wiring fault produces ignition source in main tank 1 x 10-6 8.0 8 x 10-06
Factual Information 167 Aircraft Accident Report

Boeingís fault tree analysis report contained a graphical representation of the fault

tree and a table showing the failure rates and exposure times for the fault treeís basic

elements. During its review of the fault tree analysis report, the Safety Board noted that, in

some cases, the information in the graphical representation of the fault tree did not agree

with the information in the table. For some elements, the disagreement was several orders

of a magnitude. According to Boeingís graphical representation, the probability of an

ignition event in the CWT (the top-level failure mode) was 8.45 x 10-11. Although Boeing

did not calculate the top-level failure mode probability using the table values, when Safety

Board investigators substituted the table values into the appropriate positions in the fault

tree and completed the calculations, the table values showed that the probability of an

ignition event in the CWT was 1.46 x 10-5. Table 11 shows the differences between the

data presented by Boeing in a graph and a table for of the some basic elements of its fault

tree analysis.

Table 11. Some of the basic elements of Boeingís fault tree analysis, showing the 

differences between the data presented in Boeingís graphical and tabular 

representations.

Description

Failure rate 

(graphical)

Failure 

rate 

(tabular)

Exposure 

time 

(graphical)

Exposure 

time 

(tabular)

Probability 

of failure 

(graphical)

Probability 

of failure 

(tabular)

Optimum air gap exists 4 x 10-07 4 x 10-07 0.22 8 8.8 x 10-08 3.2 x 10-06

Optimum air gap existsa 4 x 10-07 4 x 10-07 0.22 0.2 8 x 10-08 8 x 10-08

Leak through 

side-of-body rib from 

web fatigue crack

7.69 x 10-07 1 x 10-05 8 8 6.15 x 10-06 8 x 10-05

Air gap to ground exists 4 x 10-07 N/A 0.22 N/A 8 x 10-08 1 x 10-00

Scavenge pump 

deterioration or pump 

wear results in loss 

of protection

3.4 x 10-08 6.9 x 10-07 8 8 2.72 x 10-07 5.52 x 10-06

Foreign object provides 

ground in CWT N/A N/A N/A N/A 1 x 10-06 1 x 10-00

Current limit circuit 

failure (no indication) 1.1 x 10-07 8 x 10-07 8 8 8.8 x 10-07 6.4 x 10-06

Circuit breaker fails 

closed 4.76 x 10-06 4.76 x 10-06 60000 60000 2.86 x 10-01 2.48 x 10-01

Short to power on lead 

in wire bundle 2.44 x 10-08 4.07 x 10-08 8 8 1.95 x 10-07 3.26 x 10-07

Wire-to-wire short 4.07 x 10-09 2.44 x 10-07 8 8 3.26 x 10-08 1.95 x 10-06

Sufficient energy 

created to ignite fuel/air 

mixture

1 x 10-07 1 x 10-05 8 8 8 x 10-07 8 x 10-05
Factual Information 168 Aircraft Accident Report

Because of these inconsistencies in Boeingís fault tree analysis report, because of

other questions the Safety Board had regarding the detailed information used to construct

the fault tree,340 and because the fault tree did not reflect the findings of the ongoing

investigation, on June 4, 1998, the Board requested that Boeing further review and revise

the fault tree analysis report, as necessary. On July 29, 1998, Boeing responded and

indicated the following, stating that although the Board was correct in some of its points:

we [Boeing] do not believe that revising the [fault tree analysis] by incorporating

new data or changing the probability numbers will help identify new areas to

inspect or help identify the cause of the accident. We believe it would be more

productive to continue the various inspection and modification programs that are

presently underway. 

The Safety Board subsequently asked NASAís George C. Marshall Space Flight

Center (MSFC) to review Boeingís fault tree analysis report and the Boardís concerns. In a

November 25, 1998, letter, NASA MSFC personnel stated the following: 

Many of the probabilities, failure rates, and/or exposure times were much lower

than would reasonably be expected. The probability of occurrence should be

higher and/or exposure times should be longer on many of the basic

eventsÖBased on our evaluation of the tree and the information provided by the

[Safety Board], the subject fault tree analysis quantification cannot stand up to

peer review and should not be viewed as realistic. It should be noted that the logic

of the tree could not be fully evaluated since we did not have access to the

engineering drawings and schematics of the system. However, based on previous

systems experience, we would expect the tree to be constructed differently. 

NASA MSFC personnel further indicated that the exposure time numbers used in

Boeingís fault tree analysis report appeared to be overly optimistic and that the use of

more realistic numbers would result in an increased probability of a top-level failure. 

Tank sealant decays, 

exposing metal 1 x 10-07 1 x 10-06 8 8 8 x 10-07 8 x 10-06

Scavenge pump 

burnthrough results in 

ignition source

1.15 x 10-08 1 x 10-07 8 8 9.2 x 10-08 8 x 10-07

Scavenge pump 

internal friction results 

in ignition source

3.45 x 10-07 2 x 10-06 8 8 2.76 x 10-06 1.6 x 10-05

a. This event was listed in the fault tree two separate times with different values. 

340 For example, the Safety Board questioned the exposure times listed for some of the basic elements in

Boeingís fault tree analysis report. According to Boeing, most of the items listed in tables 8 and 9 had

exposure times of 0.22 to 8 hours; however, the Board notes that some of these items would likely go

undiscovered until heavy maintenance inspections (a potential exposure time of more than 2,000 hours),

whereas others might never be detected during the life of the airplane. 

Table 11. Some of the basic elements of Boeingís fault tree analysis, showing the 

differences between the data presented in Boeingís graphical and tabular 

representations.
Factual Information 169 Aircraft Accident Report

In its submission, Boeing stated that, because the fault tree analysis was developed

only to support the accident investigation, it did not go through the ìiterative review

process with the FAA normally associated with certification of a system, nor were the

underlying [failure modes and effects analyses] FMEAs developed that normally would be

the source of data to build the [fault tree analysis].î

1.16.6.5.1 Other Fault Tree Analysis Information

During its evaluation of the FMEA/fault tree process for this investigation, the

Safety Board noted several other accidents/incidents in which failures occurred that either

had not been anticipated or had not been perceived as catastrophic and, therefore, had not

been appropriately addressed in the manufacturerís fault tree analysis. For example, after a

series of accidents and incidents in the 1990s involving uncommanded rudder movements

on 737s, the Board recommended that the FAA convene an engineering test and review

board to evaluate the 737 rudder system.341 The resultant engineering test and review

board, which convened in 1999, identified catastrophic failure modes that had not been

identified through the FMEA and fault tree analyses submitted by Boeing in 1997 in

connection with certification of the 737-NG airplanes. The recent in-flight failure of a

horizontal stabilizer actuator on an Alaska Airlines MD-80 may be another example of a

failure that was not anticipated through the FMEA/fault tree analysis process.342

Further, the Safety Board observed that fault tree analyses used in noncommercial

aviation industries do not always anticipate every failure. As evidenced by the Space

Shuttle Challenger disaster, even systems that have undergone NASAís failure analysis

review can experience unexpected failure modes. In addition, the nuclear power accident

that occurred on March 28, 1979, at the Three Mile Island nuclear power plant near

Harrisburg, Pennsylvania, was attributed to a series of unanticipated failures that were,

therefore, not considered during development. According to the October 30, 1979,

ìReport of the Presidentís Commission on the Accident at Three Mile Island,î the valve

failure that initiated the accident sequence was not detected by the operators in the control

room, who relied on the indicator light indicating that the valve had closed as intended.

The operators did not heed other indications that the valve had not closed. The report

stated, ìPrior to [the Three Mile Island accident], the NRC [Nuclear Regulatory

Commission] had paid insufficient attention to [loss of coolant accidents] of this size and

the probability of their occurrence in licensing reviews.î Similarly, the NRCís January

1980 ìReport to the Commissioners and to the Publicî noted the following:

341 For additional information, see National Transportation Safety Board. 1999. Uncontrolled Descent

and Collision with Terrain, USAir Flight 427, Boeing 737-300, N513AU, near Aliquippa, Pennsylvania,

September 8, 1994. NTSB/AAR-99/01. Washington, DC. 

342 The investigation of the January 31, 2000, Alaska Airlines flight 261 accident is still ongoing (as of

the date of report publication).
Factual Information 170 Aircraft Accident Report

The greatest risk of an accident comes not from the design basis accidentsÖbut

from small loss of coolant accidents and relatively routine transients compounded

by multiple failures or human error, having a higher probability of occurring than

a large pipe break. These types of potential accident sources have, however, been

all but ignored by the NRC in the regulatory review process. 

1.16.6.6 Compensator Probe Testing at the FAAís William J. Hughes 

Technical Center 

During June/July 1999, the Safety Board conducted a series of tests at the FAAís

William J. Hughes Technical Center, in Atlantic City, New Jersey, to determine the

physical effects that the ignition and explosion of Jet A fuel vapor would have on a CWT

compensator probe assembly.343 Specifically, the tests were designed to correlate observed

damage to the compensator lead wire assemblies to the location of an ignition source in Jet

A vapor. To accomplish this goal, the tests were conducted with four different ignition

source locations. 

Video documentation of the tests showed that it was possible for a combustion

flamefront to traverse the interior of the compensator. After the tests, the test

compensators were compared with the recovered TWA flight 800 compensator. In many

instances, evidence of thermal damage (including soot deposits to lead wires and lead wire

guides and slight melting of the lead wire guide) was observed after the tests. However, in

all cases, the thermal damage observed on the test compensators was much less severe

than that observed on the recovered TWA flight 800 compensator lead wire assembly. 

1.16.6.7 Fuel Pump Tests (Motor Cavity Explosions)

The Safety Boardís review of the scavenge pump design revealed that the

scavenge pump motor was mounted on the aft surface of the rear spar (external to the

CWT, in the main landing gear wheel well) but contained fuel-cooled windings. Fuel

supplied to cool the windings passed through a flame suppression system consisting of a

small-diameter tube in series with a check valve. Further, the scavenge pump contained a

thermal fuse designed to open at less than 400° F. (The review indicated that the scavenge

pump had satisfactorily completed qualification tests during which the pump was operated

without cooling fuel until the thermal fuse opened. The Boardís review of qualification

test results and service records revealed no evidence of fuel vapor igniting inside or

outside the scavenge pump motor.) 

During its investigation of the TWA flight 800 accident, the Safety Board

conducted tests to determine the explosion-proof characteristics of an old and a new

scavenge pump with intact flame arrestor tubes and check valves. When vapors in the

motor housings of both the old and new scavenge pumps were ignited, no sparks or flames

escaped from the pump motor to ignite the explosive atmosphere in the representative

laboratory fuel tank adjacent to the pump. Further, the scavenge pumps successfully

343 For additional information, see the FAAís Summary Data Report: Investigation of the Effects of

Combustion on a Fuel Quantity Indicating System Fuel Compensator Probe, dated January 19, 2000. 
Factual Information 171 Aircraft Accident Report

contained multiple explosions that were intentionally ignited within the pump motor

cavity, preventing the ignition from transferring to the fuel tank. (Figure 37 is a diagram of

a scavenge pump.) Because the Board became aware of instances in which the flame

arrestor tube was broken or missing, investigators also conducted a series of tests to

determine whether the scavenge pump motor housing would contain explosions with only

one of the flame suppression system components (either the flame arrestor tube or the

check valve) functioning. In all tests, the scavenge pump housing contained the

explosions, and the explosive atmosphere surrounding the test pump did not ignite.

During its postaccident examination of other transport-category airplanes, the

Safety Board observed that components in the electrical connector on a scavenge pump in

a retired 747 were deteriorated. (Boeing ASBs 747-28A2194 and 747-28A0043, both

dated January 18, 1996, ìencouragedî operators to inspect for degradation in connectors.)

Investigators observed that several scavenge pump electrical connectors had been

manufactured of a material that degraded with long-term exposure to fuel. Boeing

provided the Board with a service history for the 747 scavenge pump for the 10 years

preceding the accident. When queried about the brevity of this history, Boeing indicated

that most operators did not return the scavenge pumps to the manufacturer for overhaul.

CHECK VALVE

FLAME SUPPRESSION

PASSAGE (TUBE)

(LANDING GEAR

BAY SIDE)

REAR SPAR

FORWARD

(CWT SIDE)

MOTOR

IMPELLER

Figure 37. A diagram of a scavenge pump.
Factual Information 172 Aircraft Accident Report

As a result of these findings, Boeing issued ASB 747-28A2206, dated

September 25, 1997, and AD 97-25-06, effective December 16, 1997. AD 97-25-06

contained the following summary: 

The actions specified in this AD are intended to prevent potential failures within

the electrical motor assembly of the scavenge pump, which could result in leakage

of fuel from an electrical connector into the main landing gear well, or electrical

arcing within the scavenge pump motor; those conditions could result in a fuel fire

in the wheel well. 

The AD required disconnection of the electrical connector, a one-time inspection

to identify the P/N of the electrical connector, and possible scavenge pump replacement.

The FAA subsequently received a report about damage to the internal wiring of a scavenge

pump after a replacement connector was found to have a longer backshell component.

(The longer backshell provided insufficient clearance for parts inside the pump, which

could cause arcing in the motor and fuel leakage into the wheel well.) As a result, Boeing

issued ASB 747-28A2215, dated May 14, 1998, which called for another one-time

inspection of the scavenge pump motor-impeller P/Ns, with part replacement if necessary.

The FAA made this action mandatory by releasing AD 98-14-17, effective July 23, 1998. 

1.16.6.8 Electrostatic Research 

Electrostatic hazards have been recognized for many years and have resulted in

numerous publications, including the previously discussed API 2003 and the FAA Aircraft

Lightning Protection Handbook, DOT/FAA/CT-89/222.344 The Safety Board contracted

with the AFRL at Wright Patterson Air Force Base and the Naval Research Laboratory

(NRL) to conduct research into fuels, fuel additives, and electrostatic charging of fuels.

The AFRL tests revealed that, of the various items tested, Teflon-cushioned wire clamps

(known as Adel clamps)345 were most susceptible to electrostatic charging and achieved

the highest voltage potentials. These clamps, which are used to hold tubing in the CWT,

are among those parts in the CWT that are electrically isolated (unbonded).346 (Safety

Board investigators documented more than 30 megohms of resistance between fuel tube

couplings on the test airplane, as well as between isolated metal clamps and the fuel tubes

that the clamps were mounted on, indicating inadequate bonding.) 

344 On page 198, illustration 7.27, the FAAís lightning protection guidance indicates that unbonded

clamps could present an ignition hazard if a lightning strike occurred. 

345 Teflon-cushioned, or Adel, wire clamps are used with other hardware (such as screws, washers, and

nuts) to secure electrical wires and tubing. 

346 Certification standards require that ìmajor componentsî of the powerplant installation be electrically

bonded to the other parts of the airplane. (For more information about certification requirements, see 14

CFR 25.901[b][4].) According to a June 2, 2000, letter, to the Safety Board Boeing design practices permit

parts that are less than 3 inches in any direction (including some types of clamps and connectors installed in

fuel tanks) to be electrically unbonded, presumably because such parts are not believed to have enough

capacitance to retain hazardous levels of static electricity under expected operating conditions. Boeing 747

fuel tank assembly drawings show Adel clamp installations in which the clamp does not have a grounding

strap. 
Factual Information 173 Aircraft Accident Report

Testing demonstrated that when the electrically isolated Adel clamp was exposed

to a fuel spray or drip, an electrical charge was created on the clamp. The fuel spray or

drip replicated a pressurized fuel leak. Fuel spraying or dripping onto clamps that were not

thoroughly wet (thus, not permitting the fuel to create a wet path to the ground through

antistatic additives) created higher voltage potentials than spraying or dripping onto

clamps that were thoroughly wet. However, the highest voltage potential attained in

testing was 650 volts, which would produce a discharge energy of only 0.0095 mJ. Voltage

potentials of 1,150 volts were produced in tests using fuel of a higher conductivity than

that in the accident airplaneís CWT, which could produce a discharge energy of

0.030 mJ.347

The Safety Board learned that Boeing had accumulated almost 5 J of energy on

clamps during tests that followed two explosions in Boeing 727 fuel tanks during ground

refueling at Minneapolis, Minnesota, in 1970.348 Subsequently, research conducted by

NRL personnel also found that the Adel clamp retained up to 5,800 volts (capable of

producing 0.84 mJ in a spark) when exposed to a fast drip of fuel containing commercially

unavailable additives. However, to obtain a discharge exceeding the MIE of Jet A fuel

vapor from a charged, isolated Adel clamp, the fuel used in the tests had to contain

high-conductivity additives that are no longer manufactured and not available in fuel from

JFK or the Athens airport at the time of the accident. 

1.16.6.9 Sulfide-Deposit Tests and Research

Safety Board investigators found dark deposits on FQIS fuel quantity probes,

compensators, and wiring that had been recovered from various fuel tanks in the accident

airplane and in similar locations in three other airplanes that were inspected during this

investigation (N93105, N93117, and F-BPVE [all 747s]). The Boardís examination of

several of these deposits using an SEM identified copper, silver (silver-plated copper

wiring is used in fuel tanks), and sulfur (a contaminant in jet fuel), a combination also

known as silver-sulfides.349

The Safety Board became aware that the AFRL fuel specialists had conducted

research on similar deposits in the late 1980s/early 1990s. According to the AFRL

specialists, the silver-sulfide deposits on components from the accident airplane were

similar to deposits that they had identified previously on fuel quantity probes in military

airplanes. For example, in 1990, silver-sulfide deposits were observed on a fuel quantity

probe from a USAF trainer in 1990. During electrical bench tests involving this probe,

347 Although voltage potentials of more than 5,000 volts were obtained in some laboratory tests, these

tests simulated extreme worst-case ground refueling conditions and did not use materials or fuel that would

have been present on the accident airplane. 

348 The Aviation Rulemaking Advisory Committee (ARAC) Fuel Tank Harmonization Working Group

(FTHWG) and API 2003 cite several previous fuel tank explosions that were believed to have been caused

by static electricity. 

349 Although for the purposes of this report, these deposits will be called silver-sulfide deposits, they are

sometimes referred to as ìcopper sulfide deposits,î ìsulfides,î or ìsulfidation.î According to a 1991 Boeing

engineering report, silver should never contact sulfur-containing liquids because of its susceptibility to

sulfidation. 
Factual Information 174 Aircraft Accident Report

flammable vapor ignited when voltages higher than those encountered during normal fuel

system operation were applied. The AFRL research indicated that these deposits could

break down and result in arcing when exposed to a 170-volt pulse. Further, the AFRL fuel

specialists noted that the silver-sulfide deposits were semiconductive and, thus, gradually

reduced insulation resistance between electrical connections. The AFRL report concluded

that the accumulation of silver-sulfide deposits ìis most likely the result of a long-term

degradation or corrosion processÖ[and that] as the probes age, more probe failures

[calibration errors] can be expected.î350

As a result of Safety Board requests351 for any documents pertaining to silver or

copper sulfides, on April 23, 1998, Boeing provided the Board with several documents,

including the following:

ï Boeing Materials Technology Engineering Report MS21053, dated June 17,

1988, that described the examinations of arcing damage in 18 fuel pumps. One

of the 18 fuel pumps exhibited evidence of arcing at the location of sulfide

deposits (the other 17 were too badly damaged to identify the origin of arcing). 

ï Boeing Analytical Engineering Report 2-5323-WP-91-97, dated March 20,

1991, that described Boeingís examination of sulfide deposits on a harness

from a fuel quantity indicator from the right wing tank of a 757. The fuel

quantity indicator had a documented history of irregularities after only

750 hours in service. The report noted ìit is readily observedÖthat

theÖcontaminant has migrated up into theÖconnector even after only

750 hours in service.î 

ï A Parker Bertea Aerospace memorandum, dated August 18, 1992, which

documented sulfide deposits on FQIS components removed from DC-10

airplanes.

ï Boeing Laboratory Report 9-5576-P+CA-025P, dated March 30, 1993, that

described sulfide deposits on electrical hardware from the fuel tanks of 737,

747, and DC-10 airplanes. 

ï Boeing Laboratory Report 9-5576-P+CA-025P1, dated April 29, 1993, that

also described sulfide deposits on electrical hardware from the fuel tanks of

737, 747, and DC-10 airplanes and included two additional reports (obtained

through a literature search): Silver Corrosion by Aviation Turbine Fuel, dated

September 1970, written at the Journal of the Institute of Petroleum; and

Copper and Silver Corrosion by Aviation Turbine Fuels, dated April 22, 1973,

written at the Indian Institute of Petroleum.

ï Boeing Analytical Engineering Report 9-5576-WP-97-272, dated August 5,

1997, that described electrical tests of FQIS parts containing sulfide deposits. 

350 See Slenski, George. 1990. Analysis of Trainer Aircraft Fuel Probes I. U.S. Air Force Research

Laboratory. Materials Integrity Branch, Systems Support Division, Materials Directorate. 

351 The Safety Board requested the above-mentioned information from Boeing in letters dated

December 19, 1997, and March 6, 1998. 
Factual Information 175 Aircraft Accident Report

At the Safety Boardís request, AFRL specialists at the Wright Laboratory further

examined the deposits on numerous parts from the accident airplane (including fuel

quantity probe units and an FQIS compensator fragment) and other airplanes in which

similar deposits were observed on fuel system components. 

In May 1998, the Safety Board became aware of a similar deposit found on a

Tower Air 747 CWT T347 terminal strip.352 According to Tower Air management,

maintenance personnel had been troubleshooting refueling problems (including premature

and delayed refueling shutoff) and problems with FQIS wiring when they observed the

deposits on the terminal strip and removed it. The Board subsequently requested that

several independent laboratories examine and test the deposits on the terminal strip. In

response to the Boardís request, the AFRL in Dayton, Ohio, tested the electrical resistance

of and documented the deposits; Evans-East Laboratories in Plainsboro, New Jersey,

tested the chemical properties of the deposits in greater depth; and the AFRL/University

of Dayton Research Institute (UDRI) documented the depositsí reaction to electrical

loads. In addition, tests conducted at the AFRL Electronic Failure Analysis Laboratory

and the Fuels Branch Laboratory used a portion of the Tower Air T347 terminal strip to

observe and document the depositsí breakdown in the presence of flammable jet fuel

vapor. 

The results of this research, described in reports from Evans-East Laboratories353

and the AFRL/UDRI,354 revealed that these silver-sulfide deposits accumulate over time355

as a result of exposure to jet fuel (and jet fuel vapor), which contains sulfur.356 These

deposits were typically found on FQIS wire at terminations and where damaged insulation

resulting in exposed wire conductor. Additionally, AFRL was able to ignite lighter fluid

using a sulfide deposit sample from the Tower Air T347 terminal strip. Additionally,

during tests conducted at UDRI, researchers were readily able to create numerous sulfide

deposits in a laboratory setting. According to UDRI, this research indicated that increased

temperatures enhanced the formation of silver-sulfide deposits and that fuel washing

could slow the accumulation of these deposits. 

352 Of the many FQIS components with sulfide deposits from other airplanes examined during this

investigation, the two that had been removed from the airplane because of FQIS anomalies were both 747

CWT T347 terminal strips. 

353 For additional information, see the Evans-East Laboratories reports dated April 22, 1998, and

April 24, 1998. 

354 For additional information, see the following AFRL reports: AFRL/MLSA 99-2, dated January 26,

1999; AFRL/MLSA 99-33, dated June 21, 1999; and AFRL/MLSA 99-68, dated October 18, 1999.

Additionally, an earlier AFRL report (WL/MLS 97-102) documented residues on FQIS components from

the accident airplane (see Systems Group Chairmanís Factual Report, dated November 17, 1997). 

355 UDRI research revealed that the time required for sulfide deposits to accumulate can vary

considerably and is related to factors, such as sulfur content of the fuel, temperature. 

356 Where wiring is connected to a fuel quantity probe in a 747-100 fuel tank (for example, at the

terminal block connectors), the silver-plated copper conductor of the connected wire is exposed to fuel and

fuel vapor. (The FQIS design for all 747 Classic series airplanes results in exposed conductors at terminal

blocks.) 
Factual Information 176 Aircraft Accident Report

During FAA-sponsored tests conducted in response to Safety Recommendation

A-98-37 (issued as a result of this accident), researchers at the University of Arizona also

used the method developed by UDRI to create silver-sulfide deposits in a laboratory.

During subsequent tests, these deposits served as an ignition mechanism for Jet A fuel

vapor. The results of these tests were discussed at a November 9, 1999, meeting involving

FAA, Safety Board, and industry personnel. According to University of Arizona

personnel, the formation of sulfur-containing conductive deposits from Jet A fuel on silver

wire occurred with both a.c. and d.c. current. Additional tests in this area are ongoing.357

According to a BFGoodrich representative at the November 9, 1999, meeting, on

the basis of the AFRL findings in the early 1990s, BFGoodrich had developed a means of

improving the accuracy and reliability of FQIS wiring used in military applications

through design changes that began in 1993 and involved the use of (1) nickel-plated wire,

(2) gold-plated ring connectors, (3) sealant in shrink tubes, and (4) separate inner and

outer layers of shrink tube. The BFGoodrich representative stated that previously reported

inaccuracies in the FQIS had largely resulted from current leakages through the sulfide

deposits and that the design changes resulted in a large reduction in reported FQIS

problems.358

According to Boeing personnel, Boeing uses nickel-plated, instead of

silver-plated, wiring in its newly manufactured 777 and 737-NG airplanes. However, in a

December 7, 1999, letter, Boeing indicated that ìoverall, the wholesale replacement of

FQIS bundles in the tank is not recommended.î 

1.16.6.10 Auto Ignition and Hot Surface Ignition Information

The Safety Board considered the possibility that elevated temperatures within the

CWT (resulting from engine bleed air leaks, a fire in the main landing gear wheel well, or

a fire in the air conditioning pack bay) increased the temperature of the fuel/fuel vapor or

the CWT structure/components to a sufficiently high temperature to result in ignition. This

ignition could occur as a result of ignition mechanisms known as hot surface ignition and

auto ignition. 

Hot surface ignition of Jet A fuel can result when a very high temperature surface

comes into contact with fuel or fuel vapor and is highly dependent upon many factors,

including the geometric aspects of the hot surface and the environmental conditions

affecting the fuel/fuel vapor. Existing research359 indicates that hot surface ignition of

fuels similar to Jet A can occur at temperatures ranging between 900° and 1,300° F.360

357 For information about Safety Board recommendations regarding sulfide deposits, see section 1.18.3. 

358 In a January 20, 2000, letter to the Safety Board, Boeing indicated that at least two additional

companies were manufacturing wire harnesses that have terminations resistant to sulfidation.

359 Kuchta. 

360 If an individual hot spot within the CWT were to reach temperatures between 900° and 1,300° F, it is

likely that the CWT surface in this area would also have been this hot. These temperatures are at or near the

melting temperature of the CWTís aluminum skin (aluminum softens/melts at 800° to 1,200° F, depending

on the specific aluminum alloy). 
Factual Information 177 Aircraft Accident Report

Tests conducted by Boeing indicated that strands of steel wool can be ignited by as little as

32 milliamps from a 25-volt 400-cycle a.c. source. 

Auto ignition of Jet A fuel can result when a large volume of fuel/air vapor is

sufficiently elevated in temperature for ignition to occur without direct contact with an

ignition source (such as a spark, arc, hot surface, or hot filament). Auto ignition is also

highly dependent on many factors, including the size of the heated volume and container

and environmental conditions affecting the vapor. The lowest temperature that can result

in auto ignition is called the auto ignition temperature. Existing research indicates that the

auto ignition temperature for Jet A fuel vapor at sea level is about 460° F. The auto

ignition temperature tends to increase as the altitude increases. 

Data provided by Boeing indicated that multiple engine malfunctions (concurrent

failures of the (1) engine bleed air pressure relief valve, (2) precooler cooling airflow

control valves, and (3) pylon shutoff and regulating valve would be required) could result

in bleed air temperatures above 800° F at the engines. The failures required to produce

such excessive bleed air temperatures would have resulted in an engine malfunction

warning, and the excessive bleed air temperatures would have been sensed by temperature

sensors installed inside and outside of bleed air ducts, resulting in cockpit warnings.

Further, according to engineering data from Hamilton Standard (the manufacturer of the

air conditioning system) and Boeing, the combination of failures required to produce such

elevated temperatures would have resulted in overpressure leading to an air conditioning

duct burst.361

The FAAís certification regulations regarding fuel tank temperatures (Section

25.981) state the following:

ï The highest temperature allowing a safe margin below the lowest expected

auto ignition temperature of the fuel in the fuel tanks must be determined. 

ï No temperature at any place inside any fuel tank where fuel ignition is possible

may exceed the temperature determined under paragraph (a) of this section.

This must be shown under all probable operating, failure, and malfunction

conditions of any component whose operation, failure, or malfunction could

increase the temperature inside the tank. 

1.16.6.11 Meteorite Strike Information

During the evenings after the TWA flight 800 accident, many meteorites were

observed in the northeastern United States. Therefore, the Safety Board considered the

possibility that TWA flight 800 was struck by a meteorite, which then caused the CWT

explosion. The Board consulted an expert in the study of meteorites, a professor from the

University of Pittsburghís Department of Geology and Planetary Science, to evaluate the

possibility of a meteorite hitting an airplane.362 According to the professorís testimony,

361 Examination of the recovered wreckage revealed no evidence of damage that would have resulted

from such extreme temperatures. Further, the CVR did not record any audible cockpit warning or any flight

crew discussion that would indicate that such a condition existed in the accident airplane. 
Factual Information 178 Aircraft Accident Report

knowing the frequency with which meteorites have struck cars and dwellings over the past

several decades, and based on a comparison of estimates of (1) the area of the earthís

surface represented by cars and dwellings with (2) the area of the earthís surface

represented by airplanes in the air at any given time, he calculated that a meteorite could

be expected to strike an aircraft once every 59,000 to 77,000 years. 

1.17 Operational and Management Information

1.17.1 Trans World Airlines, Inc. 

TWA holds an air carrier certificate, No. TWAA017A, and is authorized to conduct

operations in accordance with 14 CFR Part 121 as a domestic and flag air carrier. TWAís

main maintenance facility and maintenance overhaul facility is located at the Kansas City

International Airport (MCI), Kansas City, Missouri. Additionally, at the time of the

accident, TWA maintained major station maintenance bases that serviced the 747 at JFK;

LAX; MCI; STL; and San Francisco International Airport, San Francisco, California.

TWAís main operations base is located at STL. TWA conducts flight crew training at its

facilities at STL, JFK, and LAX. 

1.17.2 FAA Oversight of TWA Maintenance

The FAA principal maintenance inspector (PMI) assigned to the TWA certificate at

the time of the accident was located at the Certificate Holding District Office (CHDO) in

Kansas City, Missouri. He was hired by the FAA in July 1988 and was initially assigned as

a partial program manager (PPM) for TWAís 747, 757, and 767 fleets. The PMI was an

Assistant PMI for TWA for 2 years before he was assigned to the PMI position in 1990.

He stated that he had no certificate oversight responsibilities other than TWA and that he

had an assistant PMI and four PPMs to assist him with the oversight of the TWA

certificate. 

According to the PMI, he had adequate staffing for the oversight of TWAís

maintenance operations and did not have any concerns about the companyís maintenance

program. He stated that TWA personnel were receptive to feedback and made corrections

in a timely manner when discrepancies were identified. TWAís most recent National

Aviation Safety Inspection Program (NASIP) was conducted in September and

October 1995. The TWA NASIP involved airworthiness inspectors from Kansas City,

Missouri, who traveled to other TWA maintenance facilities to accomplish their

assignments. In addition, geographic inspectors in Atlanta, Georgia; Dallas/Fort Worth,

Texas; and Salt Lake City, Utah, assisted in ramp inspections. The Safety Boardís review

of the NASIP inspectionís airworthiness findings revealed that none of the findings

362 During his testimony at the Safety Boardís public hearing for TWA flight 800, the professor defined a

meteor as a ìvisual phenomenon caused by a body plunging through the atmosphere. You donít see the body

itself, you see the glow of compressed air and ionized gases in front of the meteorite,î and a meteorite as a

ìchunk of silicate rock or nickel-iron alloy, believed to originate in the asteroid zone of the solar system.î 
Factual Information 179 Aircraft Accident Report

pertained to 747-100 fuel systems, fuel tanks, wiring, or structural inspections. (The

accident airplane was not inspected during the NASIP.) The National Program Tracking

and Reporting Subsystem indicated, with regard to the listed airworthiness items, that

TWA corrected the items. In all cases, the Flight Standards CHDO concurred with TWAís

actions to correct the findings. 

1.18 Additional Information

1.18.1 Accident Record and History of Fuel Tank 

Fires/Explosions on Airplanes

The Safety Board has participated in the investigation of several aviation

accidents/incidents involving fuel tank explosions. According to a list prepared by the

FAA, since 1959 there have been at least 26 documented fuel tank explosions/fires in

military and civilian transport-category airplanes (including TWA flight 800).

Appendix G lists these fuel tank explosions/fires, several of which are discussed in greater

detail in this section.363

The Safety Board conducted a special investigation of the May 9, 1976, accident

involving an Iranian Air Force 747-131, as it approached Madrid, Spain, following a flight

from Iran. All 17 people on board the airplane were killed, and the airplane was destroyed.

Witnesses reported seeing lightning strike the left wing, followed by fire, explosion, and

separation of the outboard wing before the airplane crashed. Examination of the wreckage

revealed evidence of an explosion that originated in at least one of the left wing fuel tanks

near a fuel valve installation. The airplaneís fuel tanks contained a mixture of JP-4 and Jet

A fuel.364 The Boardís report noted that almost all of the electrical current of a lightning

strike would have been conducted through the aluminum structure around the ullage365 but

discussed how some energy might have entered the fuel tanks.366 Although the Boardís

report did not identify a specific point of ignition within the tank, it noted that discharges

363 Although not listed in appendix G, the Safety Board is also aware of an in-flight fuel tank explosion in

a wing tank on a U.S. Navy C-130 that occurred on January 15, 1972, near Cambridge, Maryland. The U.S.

Navyís investigation concluded that ìan arc occurred in the fuel tank airspace from eitherÖthe #10 fuel

quantity probe to an unknown groundÖ[or] from [chafing of] the coax[ial] cable in the immediate vicinity

of the #10 fuel quantity probe to an unknown ground.î

364 Although the Safety Boardís report indicated that the airplaneís fuel tanks contained a mixture of JP-4

and Jet A fuel, it is likely that the airplaneís fuel tanks actually contained a mixture of JP-4 and Jet A-1 fuel.

Jet A-1 fuel is commonly used by air carriers throughout the world, and Jet A has relatively limited

availability outside the United States. 

365 The Safety Board did not determine the probable cause of this foreign accident because it had no

statutory authority to do so. However, several hypotheses addressing the sequence of events and possible

causes of the accident were presented in the Boardís report. Further, in a February 13, 1979, memorandum,

the FAA stated, in part, ì[the Safety Boardís report indicated] that ëenergy levels required to produce a spark

will not necessarily damage metal or leave marksíÖThe enclosed photographs show the localized burned

and melted portions of the plastic sleeves of the compensator wiring harnessÖ.In view of the fact that

flammable vapors can be ignited by very low electrical energy which is less than that required to produce

any physical evidence and in consideration of the internal condition of the compensator, we suggest that a

careful review should be given of the compensator as a possible ignition source.î 
Factual Information 180 Aircraft Accident Report

could produce sufficient electrical energy to ignite the fuel/air mixture and that energy

levels required to produce a spark will not necessarily damage metal or leave marks at the

point of ignition. Upward flowing burn patterns were observed on the compensator that

the FAA considered a potential ignition source for the surge tank fire.367

The Safety Board also assisted in the investigation of the crash of Avianca

flight 203, a 727, on November 27, 1989, which killed all 107 people on board the

airplane. The airplane had departed Bogota, Colombia, about 5 minutes before the crash.

Examination of the wreckage revealed that a small bomb placed under a passenger seat,

above the CWT, had exploded.368 According to the Colombian Governmentís report

regarding this accident, the bomb explosion itself did not compromise the structural

integrity of the airplane; however, the explosion punctured the CWT and ignited the Jet A

fuel/air vapor369 in the ullage, resulting in the destruction of the airplane. 

The Safety Board also participated in the investigation of a May 11, 1990,

accident, involving a Philippine Airlines 737-300 at Ninoy Aquino International Airport,

Manila, Philippines, in which the Jet A fuel/air mixture in the CWT exploded as the

airplane was being pushed back from the gate. Of the 120 people on board the airplane,

8 were killed, and 30 were seriously injured. As a result of this accident, the Safety Board

issued Safety Recommendations A-90-100 through -103 to the FAA. In its safety

recommendation letter, the Board noted that the exact source of ignition had not been

established. However, lightning damage and damaged FQIS wires were found. The Board

stated that ìit is possible that the combination of a faulty float switch and damaged wires

providing a continuous power supply to the float switch may have caused an electrical arc

or overheating of the switch leading to the ignition of the center fuel tank vapor.î 

366 Additional information regarding the amount of energy from lightning that might reach conductive

components in fuel tanks is contained in Lightning Protection Handbooks published by the FAA and NASA. 

367 The damage was described in an undated FAA memo as follows: ìlocalized burned and melted

portions of the plastic sleeves of the compensator wiring harness where the black unshielded leadwire joins

the HI Z leadwire. This is the same area where flashover to the HI Z terminal was observed at 6.9 kilovolts

during voltage breakdown tests on new compensatorsÖ.Evidence of arcing was difficult to detect following

these tests. It may be noted in the photographs that the bottom edge of the leadwire plastic support is also

burned and that a side of the plastic sleeve around the HI Z and black leadwires is not burned. This type of

localized damage tends to indicate that a brief flame stream could have been initiated at the lug end of the HI

Z leadwire which was directed upward through the compensator as the unit is mounted in the surge tank.î 

368 Evidence of a bomb explosion included deformation of materials away from a location at the height

of the passenger seat pan, hot-gas pitting damage on multiple pieces of wreckage that formed a pattern

radiating from the same location (including into the CWT), punctures radiating from the same location, and

shrapnel. Further, according to the FBIís laboratory report, No. 91204034 S YQ YB/91207052 S YQ YB,

dated January 30, 1990, chemical analysis of a piece of wreckage from the right side of the CWT identified

the presence of RDX and PETN high explosive. These two explosives comprise about 86 percent of the

composition of SEMTEX, which is a rubberlike material manufactured by Synthesia Corporation of Semtin,

Czechoslovakia, primarily for use in mining and other civil engineering activities. According to the FBI,

SEMTEX has been used by criminal and terrorist elements in Europe since 1966. (SEMTEX was identified

as the material used in the bomb placed on Pan Am flight 103. For additional information, see section

1.11.1.2.) 

369 As indicated with regard to the Iranian Air Force 747, because Jet A fuel has limited availability

outside the United States and because Jet A-1 is commonly available outside of the United States, it is likely

that this airplaneís fuel tanks contained Jet A-1 fuel. 
Factual Information 181 Aircraft Accident Report

Safety Recommendations A-90-100 through -102 addressed the inspection of float

switch wiring on 737-300, -400, and -500 airplanes, the inspection of other wiring bundles

on all 14 CFR Part 121 airplanes, and the development and utilization of testing

techniques to detect defective float switches (with a requirement to replace all defective

units). On the basis of the FAAís responses indicating that it intended to take no action on

these recommendations, the Safety Board classified all three safety recommendations

ìClosedóUnacceptable Actionî on March 24, 1992. The Board also recommended, in

Safety Recommendation A-90-103, that the FAA conduct a detailed engineering design

review and testing of the fuel pumps used in the 737-300, -400, and -500 to verify that

overheating and interference between the rotating components of the pump and its case

will not cause a fire hazard (tests should be conducted with the jet fuel at its flash point).

The Board determined that subsequent testing and research conducted by the FAA

satisfied the intent of this recommendation and classified it ìClosedóAcceptable

Alternate Actionî on March 24, 1992. 

1.18.2 Information Regarding Electrical/Wiring Anomalies on 

Airplanes 

1.18.2.1 Accidents, Incidents, and Events Involving Electrical/Wiring 

Components 

During its investigation of the TWA flight 800 accident, the Safety Board

examined its aviation accident/incident database for records of previous accidents and

incidents involving transport-category airplanes in which the Board had used the words

ìwireî or ìwiringî in the probable cause. In addition, the Safety Board reviewed other

available records involving wiring- and/or fire/smoke-related air carrier events, including

the following: 

ï Boeing laboratory reports, SBs, and SLs; 

ï AFRL reports; 

ï Safety Board preliminary, airworthiness factual, and accident brief reports; 

ï AIR 2000 air safety reports; 

ï British Air Accidents Investigation Branch (AAIB) reports; 

ï FAA service difficulty reports (SDR);370 and 

ï a Civil Aviation Administration of China investigation report. 

The results of these reviews are discussed in this section. 

Wiring-Related Accidents/Incidents

In an October 21, 1996, fax, the Civil Aviation Authority of Singapore described

an event that occurred on October 12, 1996, in which an operator reported that arcing in a

wire bundle on a 747-200 cargo airplane had resulted in a fire at the aft bulkhead of the

forward cargo compartment about STA 1000. The airplane was undergoing maintenance

at the time of the fire, and subsequent inspection revealed damage to wire bundles W834,
Factual Information 182 Aircraft Accident Report

W846, W1524, and W370; the insulation blanket; the aft bulkhead of the forward cargo

compartment; and (possibly) the CWT sealant. The operator removed the affected

components from the airplane and shipped them to Boeing for examination and

evaluation. A December 16, 1996, letter from Boeing stated that ìX-ray microanalysis and

chemical identification of the damaged wire suggest that the insulation of the wire was

damaged and that arcing had occurred between the damaged wires or that arcing between

the damaged wires and ground had occurred.î 

The Safety Boardís database research revealed 15 Board-investigated air carrier

accidents/incidents that occurred between 1983 and 1999 in which wiring was a factor in

the probable cause. These events occurred on a wide variety of airplane

types/manufacturers (including a 727, two 737s, a 747, a 767, a DC-8, a DC-9, an MD-88,

three DC-10s, an L-1011, an A320, a DeHavilland DHC-8, and an ATR) and various air

carriers. A review of these records indicated the following anomalous electrical/wiring

conditions: 

ï On May 27, 1983, near Kansas City, Missouri, a wire bundle behind the

L-1011ís flight engineerís panel electrically short-circuitedóthe flight

engineer extinguished the resultant fire with a portable fire extinguisher.

Smoke in the cockpit abated naturally.

ï On October 7, 1983, at Juneau, Alaska, a taxi light switch on the departing

727-247 series airplane failed and burned the insulation on an adjacent wire. 

ï On May 17, 1984, near Minneapolis, Minnesota, the first officer on a DC-10

observed sparks and smoke from the left-side panel. Cockpit lights were turned

off, and the smoke dissipated. Subsequent inspection revealed that a wire

bundle located between the center of the instrument panel and a fluorescent

light beneath the first officerís glare shield exhibited evidence of fire damage.

This wire bundle was stretched tightly across a sharp metal edge on the side of

the fluorescent light. Insulation and a fabric sleeve around the wires were

burned, exposing the wires. Investigation revealed that the wire bundle was

removed and reinstalled 4 days before the event.

370 Although the FAA collects a significant amount of data about mechanical failures through its SDR

program, the Safety Board has noted that these data are often incomplete. In a June 3, 1993, letter to the

FAA, the Board stated that ìthe SDR program is often used during aviation accident/incident investigations

to research the history of aircraft failures, malfunctions, and defects. However, attempts to effectively use

the SDR database in recent Safety Board investigations have revealed that the current program is incomplete

and of limited value in identifying service defect histories because many reportable service difficulties are

not reported to the FAA.î In the same letter, the Board issued Safety Recommendations A-93-61 through -63

to the FAA, which proposed changes to improve the SDR program. On August 4, 1995, the FAA issued

NPRM 95-12, ìOperational and Structural Difficulty Reports,î proposing to revise the reporting

requirements, but no final rule was ever issued. In subsequent investigations, the Board continued to find

incomplete information in searches of the SDR database and on January 9, 1998, issued Safety

Recommendation A-97-125, which asked the FAA to modify the SDR reporting system so that it would

contain more complete and accurate information about component failures. On April 15, 1999, the FAA

agreed that improved reporting was needed and published a supplemental NPRM to update the proposed

changes. On October 26, 1999, Safety Recommendations A-93-61 and A-97-125 were classified ìOpenó

Acceptable Response,î pending issuance of the final rule. The FAA issued the final rule on September 15,

2000. 
Factual Information 183 Aircraft Accident Report

ï On October 5, 1986, at Seattle, Washington, the internal power feeder cables

on a DC-10 short-circuited and melted during an attempt to apply ground

power to the airplane. According to the Safety Boardís report, the investigation

revealed ìpoor workmanshipî in a modification installation.

ï On July 12, 1987, near Boston, Massachusetts, faulty wiring in the throttle

quadrant of the 767 airplane resulted in an uncommanded go-around. 

ï On March 30, 1988, at Boston, Massachusetts, a loosened battery ground

connection on a DC-10 airplane resulted in an electrical arc and ignition of

pooled fuel in the center accessory compartment. There was fire damage to the

compartment, wires, and insulation blanket. 

ï On April 15, 1988, near Seattle, Washington, an electrical wire in the DHC-8

airplane chafed against a hydraulic line and arced, resulting in a hydraulic fluid

leak, a subsequent fire, and hydraulic system failure.

ï On December 30, 1989, at Tucson, Arizona, a 737 experienced an in-flight fire

and loss of hydraulic power. Inspection revealed that an electrical wire chafed

against a hydraulic line; the wire arced, resulting in a hole in the hydraulic line

and a hydraulic fluid leak. Subsequent arcing ignited the leaking fluid,

resulting in a fire and loss of hydraulic power. 

ï On November 13, 1991, near Toledo, Ohio, damaged wires in the DC-8

airplane wire bundle resulted in a faulty indication that the main cargo door

was closed and locked. The main cargo door subsequently opened in flight. 

ï On November 20, 1991, while departing Atlanta, Georgia, a passenger on

board the MD-88 observed fire coming from the right-side overhead

compartment. Subsequent inspection revealed that a damaged cannon plug

connector adjacent to a cabin overhead light assembly resulted in an electrical

short circuit. 

ï On January 3, 1992, near Denver, Colorado, the flight crew on a 737 noted an

ìelectrical-typeî burning odor, followed by a loud arcing sound and a small

fire at the flight deck aft overhead panel. The flight crew reported that it easily

extinguished the fire. Subsequent inspection revealed that a wire bundle chafed

against the cockpit door frame and short-circuited, resulting in electrical arcing

and the small fire. This wire bundle was not shielded or anchored to protect

against chafing, as was the other wire bundle routed through this compartment. 

ï On February 25, 1994, the flight crew on the ATR observed smoke in the

cockpit during descent near Mosinee, Wisconsin. Subsequent inspection

revealed that a push-button selector indicator switch in the cockpit had melted

and resolidified, and adjacent wiring exhibited evidence of heat damage.

Records indicated previous malfunctions associated with water/fluid

contamination near the switches, and the manufacturer had issued two SBs

addressing the problem. 
Factual Information 184 Aircraft Accident Report

ï On March 18, 1996, near Detroit, Michigan, the flight crew on the Airbus

A320 airplane received a series of elevator/pitch-related computer fault

messages. Subsequent inspection revealed an intermittent fault in the elevator

servo controller electrovalve coil and a defective diode in an electronic module

for one of the stabilizer actuators. 

ï On July 5, 1997, near Flint, Michigan, the DC-9 flight crew observed smoke

beneath the instrument panel glareshield. Subsequent inspection revealed

burned wiring and a charred position light switch behind the cockpit

instrument panel/glareshield. Further examination revealed an electrical short

in the left wing tip oscillating position light autotransformer. 

ï On May 18, 1998, near Minneapolis, Minnesota, the 747 flight crew

experienced several uncommanded ìkicksî of the yaw damper while

troubleshooting a fire warning indication. The flight crew disengaged the yaw

damper but continued to experience uncommanded ìkicks,î although they

were less severe. Subsequent inspection revealed that a VHF transceiver coax

connector located near the yaw damper control box was not properly shielded,

resulting in EMI with the yaw damper system. 

FAA Service Difficulty Report Data Regarding Wiring and/or 

Fire/Smoke-Related Events

The Safety Board also searched the FAAís SDR database for wiring and/or

fire/smoke-related events involving 747s and found 21 events.371 These events are

summarized in the SDRs as follows:

ï Electrical fire in a galley. Investigation revealed that the source of the fire was

damaged wires in a coffeemaker. (FAA control No. 00013, December 21,

1986.) 

ï Electrical fire started in a passenger entertainment control box near the R4

door. (FAA control No. 00030, June 12, 1987.)

ï Thirteen burned wires in a 6- by 8-inch area at the cabin floor line were found

at STA 1200. Circuit breaker for pallet drive wheels tripped. (FAA control

No. 00029, December 2, 1988.) 

ï Electrical fire from an electrical short circuit in a galley. (FAA control

No. 00119, February 2, 1990.)

ï Electrical fire below aft galley personnel elevator. (FAA control No. 0008,

October 29, 1990.) 

ï Electrical fire in the upper control module of the first-class galley. Found a

split capacitor in the oven controller. (FAA control No. 00135, December 21,

1990.)

371 This list does not include events involving light ballasts, failures to light sockets, reseated electrical

connectors that cleared faults, and secondary damage to wires caused by failure of other components. 
Factual Information 185 Aircraft Accident Report

ï The No. 2 leading edge device position switch was short-circuited to the

ground, resulting in a malfunction indication (amber light) in the cockpit.

(FAA control No. 0015, March 15, 1996.)

ï R4 door emergency lights were inoperative because of loose wiring on a

cannon plug at the battery charging unit. (FAA control No. 00466, April 5,

1996.)

ï R4 and L2 emergency lights were inoperative. Maintenance actions included

repairs to the emergency light pin and wiring. (FAA control No. 00199,

June 20, 1996.)372

ï The No. 3 engine fire loop wires were worn through at the hydraulic box

forward bracket. (FAA control No. 00263, August 8, 1996.)

ï During a daily check, maintenance personnel found zone E right-side

emergency path light inoperative. Replaced fuse and repaired chafed wire.

(FAA control No. 00572, August 15, 1996.) 

ï Left main landing gear microswitch wires were chafed, resulting in cockpit

indication that the left landing gear was down and locked when the landing

gear was retracted. (FAA control No. 00662, August 15, 1996.) 

ï During a daily check, maintenance personnel found two lavatory smoke

detectors inoperative. Found and repaired broken wiring. (FAA control

No. 00621, August 15, 1996.) 

ï Floor emergency lights inoperative on the left side of the cabin at seat rows 51

to 66 and on the right side of the cabin at seat rows 35 to 43. Secured

emergency light wiring. (FAA control No. 00112, August 29, 1996.) 

ï During a daily check, maintenance personnel found 5L emergency light

inoperative. Repaired wiring. (FAA control No. 00760, August 29, 1996.) 

ï A wire short circuited to the ground and resulted in inoperative upper deck

floor evacuation lights. (FAA control No. 00525, September 5, 1996.) 

ï Electrical arcing sounds and sparks were observed near the overhead passenger

service unit at seat 4B. Examination revealed that the wiring harness that runs

parallel with the sidewall light assembly above 4A and B was burned. (FAA

control No. 00231, September 5, 1996.) 

ï During a daily check, maintenance personnel found No. 2L (main entry door)

upper light inoperative. Found chafed wire at bustle light connector. (FAA

control No. 00405, September 12, 1996.) 

ï Disconnected wiring resulted in inoperative floor evacuation path emergency

lights from cabin seat row 3 to 27. Maintenance personnel reconnected wiring

and installed light assembly. (FAA control No. 00633, September 19, 1996.) 

372 Maintenance records indicated that this SDR pertained to repairs performed on the accident airplane.

The repaired wire was part of a bundle that branched off from a larger bundle that included CWT and left

wing FQIS wires leading to the upper deck AIDS unit and also contained high-voltage wiring for lighting.

Because condensation is a common problem in transport airplanes, the 747 design includes drip shields to

protect critical elements (including cockpit wiring) from condensation. 
Factual Information 186 Aircraft Accident Report

ï Electrical arcing and smoke were observed from work light wiring for a galley

coffeemaker. (FAA control No. 00668, September 26, 1996.) 

ï Floor evacuation path light wiring short-circuited near cabin seat 50H. (FAA

control No. 00747, October 10, 1996.) 

1.18.2.2 Previous Safety Board Safety Recommendations Regarding 

Electrical Components

The Safety Board reviewed its safety recommendation database for

recommendations that involved electrical components by searching for key words, such as

ìbreaker,î ìcircuit,î ìcircuitry,î ìelectrical,î ìwire,î and ìwiring.î The safety

recommendations identified in the Boardís review were evaluated, and several

nonpertinent recommendations were eliminated.373 This section discusses the pertinent

safety recommendations. 

On January 12, 1967, the Safety Board issued Safety Recommendation A-67-003

to the FAA, which addressed electrical cockpit heaters and related circuitry in

Convair 340, 440, and 580 model airplanes. The safety recommendation letter indicated

that ìwhile examining the aircraft, loose nuts, screws, drill shavings, and trash were found

inside the circuit breaker panel compartment.î The FAA responded in a February 2, 1967,

letter that ìto preclude accumulation of trash and unwanted items in the circuit breaker

panel compartment, Allegheny Airlines is in the process of issuing an inspection card

calling for vacuuming this compartment each 300 hours.î On January 1, 1975, the Safety

Board classified Safety Recommendation A-67-003 ìClosedóAcceptable Action.î 

On June 2, 1983, an in-flight fire occurred on board a McDonnell Douglas

DC-9-32, operated as Air Canada flight 797, and, after an emergency landing, the cabin

interior continued to burn. Five crewmembers and 19 passengers were able to evacuate the

burning airplane; the remaining 23 passengers died in the fire. On July 19, 1983, the

Safety Board issued Safety Recommendation A-83-047, which recommended that the

FAA require the inspection and periodic reinspection of the lavatory flush pump motors

and associated wiring; Safety Recommendation A-83-048, which recommended that the

FAA develop a procedure to inspect lavatory flush pump motors; and Safety

Recommendation A-83-049, which recommended that the FAA require bulletins for

frequent removal of waste from lavatories, ìwhich gives sufficient emphasis to areas

susceptible to the accumulation of fluids in the vicinity of wire harnesses and other

electrical components which can cause corrosion.î On June 12, 1984, the Safety Board

classified Safety Recommendation A-83-049 ìClosedóAcceptable Action.î On

November 17, 1986, the Safety Board classified Safety Recommendations A-83-047

and -048 ìClosedóAcceptable Alternate Action.î374

On June 27, 1988, the Safety Board issued Safety Recommendation A-88-064,

which recommended that the FAA conduct a directed safety investigation to determine the

373 Nonpertinent safety recommendations included those in which the electrical component key words

were only indirectly referenced in the discussion of other systems (for example, circuit breakers were

referenced in a safety recommendation regarding CVRs). 
Factual Information 187 Aircraft Accident Report

reliability of circuit breakers and the mechanisms by which failures internal to the circuit

breakers can disable operating systems and to identify appropriate corrective actions, as

necessary. The FAA agreed with the intent of the recommendation and reported that an

examination had been made. On October 24, 1989, the Safety Board classified Safety

Recommendation A-88-064 ìClosedóAcceptable Alternate Action.î 

On August 8, 1990, the Safety Board issued Safety Recommendation A-90-102,

which recommended that the FAA issue an AD ìapplicable to all 14 CFR Part 121

airplanes to require, at the next scheduled maintenance inspection, an inspection of the

wires in wire bundles in the wings where additional wiring has been added since the

airplane was manufactured. The inspection should be directed to the determination of

insulation damage where the bundle is under clamps and inside vapor seals and pressure

seals.î The FAA refused to issue the recommended ADs and cited the Boeing 737

Maintenance Planning Document, which recommended a zonal visual inspection of the

interior CWT that would include wires and wire bundles at every 7C check

(approximately 21,000 flight hours). On March 24, 1992, the Safety Board classified

Safety Recommendation A-90-102 ìClosedóUnacceptable Action.î 

On August 14, 1991, the Safety Board issued Safety Recommendation A-91-070,

which recommended that the FAA require ìspecific quality control and inspection

procedures for wire bundle installations on transport-category aircraft to verify proper

bend radii, chafe protection, and routing practices by aircraft manufacturers during

fabrication and by airlines during maintenance operations that expose wire bundles.î The

FAA agreed with the intent of this safety recommendation and required the recommended

quality control and inspection procedures. On February 10, 1994, the Safety Board

classified Safety Recommendation A-91-070 ìClosedóAcceptable Action.î 

Also on August 14, 1991, the Safety Board issued Safety Recommendation

A-91-071, which recommended that the FAA ìnotifyÖ[PMIs] and operators of

transport-category aircraft of the fire hazard posed by accumulations of lint and other

debris on wire bundles.î The FAA subsequently issued Handbook Bulletin 91-15, ìOrigin

& Propagation of Inaccessible Aircraft Fire Under In-Flight Airflow Conditions,î and

recommended that PMIs ensure that operators of transport-category airplanes clean wiring

in inaccessible areas. On October 20, 1992, the Safety Board classified Safety

Recommendation A-91-071 ìClosedóAcceptable Action.î 

On April 9, 1992, the Safety Board issued Safety Recommendation A-92-021,

which recommended that the FAA ìrequire that the electrical actuating systems for

nonplug cargo doors on transport-category aircraft provide for the removal of all electrical

power from circuits on the door after closing (except for any indicating circuit power

necessary to provide positive indication that the door is properly latched and locked) to

eliminate the possibility of uncommanded actuator movements caused by wiring short

374 On May 10, 1985, lavatory fluid was found to have contaminated the wire bundles near improper

(deep) hot-stamp markings, resulting in an in-flight fire aboard a Monarch Airlines 757. Additionally, on

October 22, 1995, lavatory fluid-contaminated electrical components (specifically a connector in the yaw

damper circuit) resulted in uncommanded roll/yaw oscillations in a 737. 
Factual Information 188 Aircraft Accident Report

circuits.î The FAA agreed with the intent of this safety recommendation and required that

all such electrical power be removed before taxi. On August 10, 1993, the Safety Board

classified Safety Recommendation A-92-021 ìClosedóAcceptable Alternate Action.î 

1.18.2.3 Inspection of Wiring in Several Transport-Category Airplanes

Postaccident inspection of recovered wiring from the accident airplane and other

transport-category airplanes revealed numerous notable conditions in both older and

newer transport-category airplanes, including damage to wire insulations, fluid stains on

wires, and metal shavings resting on and between wires in wire bundles. The Safety Board

is aware of several prior incidents that resulted from short circuits apparently caused by

contaminants. For example, after a January 9, 1998, short-circuit incident aboard a United

Airlines 767, copper conductors from wires within the damaged wire bundle were found

with melted aluminum (consistent with metal shavings found immediately beneath the

wire bundle), indicating that a metal shaving acted as a bridge to form a short circuit.

Another example is the October 22, 1995, incident involving roll/yaw oscillations on a

737 near Manchester, England. The incident was attributed to a short circuit in the rudder

circuitry resulting from lavatory fluid bridging the pins within an electrical conductor.

Additional similar incidents are described in the Systems Group Chairmanís Factual

ReportóAddendum for Aircraft Wire Inspections and Historical Reports, including the

following: a July 19, 1997, incident involving a Lufthansa 747 (after circuit breakers

opened while on approach to JFK, drill shavings were found in an area of burned and

damaged wiring in the front cargo hold) and a May 10, 1985, incident involving a

Monarch Airlines 757 (after generators tripped and smoke appeared in flight, lavatory

fluid was found on a burned area of a wire bundle, and insulation was found damaged by

out-of-tolerance hot-stamp markings).375

In connection with the TWA flight 800 investigation, the Safety Board examined

the wiring in 25 transport-category airplanes that had been maintained by various air

carriers. The inspected airplanes included 18 Boeing airplanes (13 747-100, -200, and

-300 ìClassicî series airplanes;376 a 747-400; 2 737s; a 757; and a 767), 4 Douglas

airplanes (a DC-9-30, an MD-90, and 2 MD-11s), 2 Lockheed L-1011s, and an Airbus

A300. The airplanes examined ranged in age from new (not yet delivered to the operator)

to 28 years old (retired by the operator) and in total flight time from 4 to 102,712 hours. 

The Safety Board used manufacturer and air carrier wiring diagrams as references

during its inspections of these airplanes. Although different airplane types and series were

inspected, the same general areas were inspected in each airplane when accessible. These

areas included the main electrical equipment center,377 circuit breaker panels, the area

behind the flight engineerís panel and forward of the cockpit escape door, the area near the

375 Additional cases in which contamination was found are discussed later in this section and in

sections 1.18.2.1 and 1.18.2.2.

376 The airplanes examined were from various sources. The investigative group agreed that the general

condition (extent of contamination and workmanship) of the wiring observed in airplanes of a given operator

was similar to that of other operatorsí airplanes with similar flight times. 
Factual Information 189 Aircraft Accident Report

water bottles in the forward cargo compartment, and the pressure seal-to-wing wire

route.378 Figure 38 shows the areas typically inspected on 747 airplanes. 

Metal debris was observed on or near the wire bundles of all airplanes except for

one new 737, which was only examined by investigators for about 10 minutes. A

9-month-old airplane had light lint ìfuzzî beginning to accumulate on wires behind an

avionics rack. Additionally, accumulations of foreign materials were observed on and

between wires/wire bundles in many of the airplanes in many locations. The foreign

materials included lint (fiber and dust particles), metal shavings, hardware (washers,

screws, rivets [drilled and undrilled]), structural corrosion preventive compound, blue and

brown fluid stains, paint, and pieces of paper. Wire insulation was damaged and/or cut

where metal debris was found between individual wires within a wire bundle. In some

cases, the core conductors of the wires were exposed to the grounded shielding of other

wires. Wire bundles in wheel wells were typically black and oily, and fluid was found on

wires located near hydraulic actuators in wings; grease was found on wires near flap

actuators and on landing gear struts. Wire bundles were found adhered into solid, stiff

masses with additional wires strapped to the outside of the existing bundle. The

inspections also revealed crumbled rubber cushions in clamps, cracked O-rings, kinked

wires, and wires with cracked and/or chafed insulation. 

Further, investigators identified several instances in which the wires and/or wire

bundle routing, clamping, and/or tension in various series 747 airplanes did not appear to

be in accordance with Boeingís 747 manuals or PIs. For example, in some airplanes,

parallel small wire bundles were tied into larger wire bundles differently than shown in the

PIs. Additionally, in some airplanes, wires were stretched tightly between clamps and

rubbed/chafed against clamps, other brackets, and around bends. This rubbing/chafing

often resulted in abraded wire insulation. In one airplane, three blackened rub marks were

found on two wire bundles. In one 747, a set of twisted wires in a woven jacket (similar to

that used for FQIS wires) had chafed through the jacket, wire topcoat, and into the base of

the insulation. Further, in other inspected airplanes, wires/wire bundles that were supposed

to be segregated in separate parallel bundles (according to Boeing 747 PIs) were in contact

with each other. 

377 The term ìmain electrical equipment centerî is used in Boeingís 747 Systems Handbook to refer to a

compartment located aft of the nose landing gear bay that contains avionics and electrical distribution

components. Although Boeing describes similar compartments in its airplanes by various terms, they are

generally referred to as ìE/E bays.î 

378 For detailed information regarding the individual airplane inspections, see Systems Group

Chairmanís Factual ReportóAddendum for Aircraft Wire Inspections and Historical Reports, dated July 28,

1999. 
Figure 38. Diagram showing the areas typically inspected on 747 airplanes.

FLIGHT ENGINEERSTATION (RIGHT)

COCKPIT ESCAPEDOOR (RIGHT) STA 920

STA 100

WATER BOTON FORWAR

PRESSURE SWING WIRE R

FORWARD CARGOCOMPARTMENT

AVIONICS

STA 400

NOSE LANDINGGEAR COMPARTMENT
Factual Information 191 Aircraft Accident Report

In one airplane (its last known flight was in January 1997), wires in the raceway

over the CWT were not bundled (as indicated by Boeingís PIs) but were randomly laid in

a tray, such that individual wires crossed from one side of the tray to the other.

Higher-voltage generator and galley power cables were clamped to the adjacent structure

(not routed in the tray). Inspection of the inside of this airplaneís CWT revealed the

following: the scavenge pump inlet screen contained a nearly 1-inch-long uninsulated

wire fragment, a vertically routed FQIS wire was chafed where it contacted a horizontal

structural flange, and a horizontally routed FQIS wire was chafed where it contacted a

vertical row of rivets between two supporting clamps. Examination of this airplaneís CWT

fuel quantity probes by the AFRL experts revealed sulfide deposits on the bare wire ends

and terminal sleeves, and some of the wires were loose under the clamps; two of the wires

exhibited striation marks that resembled jaw marks of a wire stripping tool. 

Five of the 25 airplanes inspected by Safety Board and other investigators during

this investigation exhibited signs of fire and heat damage, as evidenced by the following: 

ï In the 767 E/E bay near the galley chiller unit, investigators observed melted

and resolidified wires, charred insulation and burned foam, black and crispy

wire insulation adjacent to a woven fabric sleeve around a small wire bundle,

and copper balls consistent with arcing on wires more than 1 inch from wires

exposed through melted holes in wire insulation. Additionally, an AAIB

inspector was in the compartment when the mechanic applied electrical power

to the airplane during the inspection, and the wiring aft of the galley chiller

began to arc.379 (Investigation revealed that the galley chiller had been replaced

the previous day. Water was found on a thermal insulation blanket beneath the

galley chiller and wire bundle. Metal shavings, coins, screws, copper wire,

locking wire, plastic cable ties, and dust/dirt were also found on or beneath

wire bundles.) 

ï In one of the L-1011ís mid-electronics-service-center,380 two wire bundles of

about 3-inch diameter were found burned. Examination of the area revealed

blue-fluid stains and dust accumulations on wire bundles and generator feeder

cables. In the burned area of the bundles, melted and resolidified copper

provided evidence of wire-to-wire arcing; there was no evidence of arcing to

other structure or areas.381

379 After this incident, the FAA issued AD 98-07-26, which required 767 operators to inspect the

airplanesí wiring and to wrap the wires for additional protection. The FAA indicated that its examinations of

150 other 767 airplanes revealed similar damage to wiring near the galley chillers in 13 of the airplanes. 

380 Lockheed uses the term ìmid-electronics-service-centerî to describe an area located forward of the

forward wing spar, beneath the aft edge of the 2L lavatory in the cabin. As noted previously, Boeing refers to

similar areas in its airplanes as ìE/E bays.î 

381 After this damage was found, the operator inspected the remainder of its L-1011 fleet (35 airplanes).

The operator reported the following findings: 5 airplanes exhibited evidence of blue-fluid contamination of

wires in the inspection area; 11 airplanes exhibited ìsome signs of feeder to bundle rubbingî in the

inspection area; 17 airplanes ìdemonstrated excessive dirt and dustî in the inspection area; and 13 airplanes

had metal shavings and/or paper debris in the inspection area. Further, the operator reported that it was

aware of three or four previous occurrences involving blue-fluid contamination of wire bundles below the

midlavatories on its L-1011s. 
Factual Information 192 Aircraft Accident Report

ï Examination of the video system control unit (VSCU) near one of the MD-11ís

R2 door revealed melted and resolidified plastic and a low-resistance path

between pins in the cannon plug electrical connector that had been attached to

the unit. Teardown of the VSCU revealed liquid stains on the top and bottom

surfaces of internal circuit boards. The manufacturer stated that the circuit

connected 115-volt a.c. power to circuits designed for 28 volts or less. 

ï Beneath the other MD-11ís floorboards (between STAs 1661 and 1681), a

metalized mylar insulation blanket (a 60 inch X 20 inch area) was

burned/partially burned. Detailed examination of this area revealed that a

wiring harness containing 20-gauge insulated wire was routed across the frame

at STA 1681. One wire was separated, and seven other wires exhibited

damaged and chafed insulation where they contacted the frame. A 1 1/4-inch

hole in the insulation blanket was located beneath the chafed wire bundle. 

ï Behind the circuit breaker panel to the right of the flight engineerís station on

the other L-1011, an Adel clamp and circuit breaker exhibited evidence of

thermal damage (blackening). One circuit breaker (CB 3F1) was found

popped, and the circuit breakerís post appeared to have arced to an Adel wiring

clamp that held a bundle of wires. Maintenance personnel noted that the wire

stand-off associated with this Adel clamp was about 1/8 inch long. They

indicated that the wire bundle worked its way close to the circuit breaker lug

until the Adel clamp contacted the lug and shorted. 

1.18.2.4 Boeing Service Letter Regarding Inspection of Wiring on High-Time 

747 Airplanes (747-SL-20-048, Dated January 25, 1995) 

Boeing indicated that after being ìrequested by operators to provide guidance on

areas of wiring that warrant special attention on high time 747 airplanes,î it inspected

several high-time 747 airplanes. On the basis of these inspections, on January 25, 1995,

Boeing issued SL 747-SL-20-048, which stated that ìoverall, Boeing has found that most

airplane wiring exceeds the economic design goal of the airplane.î382 Further, the SL

stated that Boeing ìbelieve[d] that the wiring on high time 747 airplanes is holding up

exceptionally well,î and noted that ì[w]iring damage is hard to predict but some areas of

wiring experience degradation more frequently.î Boeing indicated that it considered ìthe

principal causes of wiring degradationî to be vibration (and vibration in conjunction with

other factors), maintenance (proper and improper),383 indirect damage (damage resulting

382 Boeing indicated that the 747ís economic design goal was 50,000 flight hours; however, it noted that

many in-service 747s significantly exceed (in some cases, more than double) this flight time. (According to

Airclaims, as of November 30, 1999, 469 of the 747s operated worldwide had more than 50,000 hours,

including 14 with more than 100,000 hours. Airclaims data indicated that 154 of the 747s operated at that

time by U.S.-based operators had more than 50,000 hours, including 3 with more than 100,000 hours.)

Although Boeing indicated that the 747ís economic design goal was 50,000 flight hours, Boeingís design

documentation for the electrical system indicated that the design goal of electrical system components was

ì30,000 hours of continuous trouble-free operation.î 

383 SL 747-SL-20-048 noted that although improper maintenance can contribute to long-term wiring

problems and wiring degradation, even proper maintenance can disturb wiring, resulting in increased

degradation. According to Boeing, ìas wiring and components become more brittle with age this effect

becomes more pronounced.î
Factual Information 193 Aircraft Accident Report

from events not directly related to wiring, such as pneumatic duct ruptures), chemical

contamination,384 and heat.385

SL 747-SL-20-048 stated that operators should be particularly attentive to the

following areas during wiring inspections: 

ï Clamping pointsódamaged clamps, clamp cushion migration, or improper

clamp installations can aggravate wire chafing. 

ï Connectorsóworn seals, missing seal plugs or dummy contacts, or lateral

pressure on connector grommets can compromise connector integrity and

allow contamination to enter the connector, resulting in corrosion or grommet

degradation. 

ï Terminationsóterminal lugs and wire splices are susceptible to mechanical

damage, corrosion, heat damage, and chemical contamination. Also, the

buildup and nut torque on large-gauge wire studs are critical to their

performance. 

ï Backshells386ówires may break at backshells because of excessive flexing,

static pressure, or missing build up. 

ï Sleeves and conduitsódamage to sleeving and conduits, if not corrected, will

often lead to wire damage. 

SL 747-SL-20-048 also stated that a special inspection of high-time 747 airplane

wiring should include the following areas: wings, engine and nacelle, landing gear,

electrical panels, and power feeders. Boeing noted that the wings, engine and nacelle

areas, and landing gear are high vibration areas. The SL further stated that ìthe wing

leading edge and pylons areÖdifficult environment[s] for wiring installationsî and, in

attachment 2 to 747-SL-20-048, Boeing indicated that it could ìsee the virtue of a wire by

wire inspectionî of these areas. The attachment described such an inspection as follows: 

Wiring is released, cleaned, bundle clampsÖand bundle ties removed. However,

environmentally sealed and potted installations and wiring in conduit should be

left undisturbed unless damage is suspected. Wiring should be closely inspected

for damage from vibration and exposure to environmental damage. Wire bundles

are separated to allow inspection of individual wires. 

384 SL 747-SL-20-048 stated that ìcertain chemicals commonly found in an air carrier operating

environment (for example, hydraulic, cleaning, corrosion-inhibiting, and/or deicing fluids) can result in

wiring component degradation.î

385 According to Boeing, wiring that was not designed for high temperatures will degrade more rapidly

when exposed to heat. SL 747-SL-20-048 stated that exposure to even low levels of heat can degrade wiring

over long periods of time. 

386 A backshell is a mechanical component that provides protection and support for wires attaching to an

electrical connector, relieving strain from the electrical pins to which the wires attach. 
Factual Information 194 Aircraft Accident Report

1.18.2.5 Wiring Inspection Methods 

Currently, aircraft wiring is usually visually inspected during an air carrierís

detailed inspection of the area through which it is routed.387 This method of wire

inspection requires a visible defect to develop before detection is possible. However, as

previously stated, a large portion of an airplaneís electrical wiring is not readily visible

(and, therefore, not readily inspectable) because it is bundled with dozens of other wires

(of which, only the portions of the wires on the surface of the bundles would be visible) or

blocked from view by other structures or components (including other wire bundles,

airframe structure, and avionics or other system components). Other obstructions to visual

inspection of wires include protective sleeves installed on wires, clamps, potting material

or sealant, and limited viewing angles because of confined inspection areas. 

Although many wire defects may be difficult or impossible to detect through visual

inspection alone, automated test equipment (ATE) inspection systems are available to

supplement visual inspections. These systems include electrical continuity or resistance

tests, insulation resistance and capacitance tests, and time-domain reflectometry (TDR).388

According to the manufacturers, ATE inspection systems would be capable of quick,

thorough, and reliable inspections of electrical system components, such as wiring and

connectors, on commercial airplanes. ATE inspection systems can be connected to an

airplaneís electrical system through electrical connectors to perform a series of

computerized inspections of every wire within a bundleóeven those that would not be

visible to the human eye. ATE inspection systems could be used to establish baseline

properties of individual wires within a wire bundle and to detect any differences in the

properties of those wires (possible indications of the presence of short circuits, chafing,

abrasions, cuts, or cracks in the insulation) during subsequent inspections. Using ATE

inspection systems would allow air carriers to evaluate the condition of electrical wiring

and detect defects before they become visually apparent and cause an electrical

malfunction. Further, the location of the degradation could be determined from these test

sets, which would allow specific repairs to be performed before failures (such as electrical

arcing) could occur (known as proactive inspection). 

However, ATE inspection systems have some limitations and disadvantages,

including equipment cost, size, and weight; the need for specially trained test equipment

personnel and numerous cables long enough to attach the ATE to the airplane at the

furthest ends of individual circuits; unique electrical connectors that must mate with

connectors in the airplanes to be tested;389 and preprogrammed software requirements.

According to observers, the time required to set up the test equipment, then disconnect the

test equipment and reinstall airplane components that had been removed for access to

connectors, exceeded the time it took to run the test. 

387 For a description of TWAís general inspection procedures, see section 1.6.3.1.1. 

388 TDR is an ATE inspection technique that involves injecting small electrical pulses into each wire and

then measuring the reflected energy. 

389 There are numerous electrical connectors available to manufacturers and operators, and the specific

connectors used may vary from airplane to airplane and within an airplane, making it difficult to match with

ATE connectors. 
Factual Information 195 Aircraft Accident Report

The Safety Board has learned that since the TWA flight 800 accident, the FAA,

U.S. Navy, and USAF have (independently and jointly) conducted research into ATE

aircraft wiring inspection systems/techniques. These research efforts include the

following: 

ï Through the Aging Transport Systems Rulemaking Advisory Committee

(ATSRAC),390 the FAA has tested/evaluated ATE systems from several

manufacturers at the FAAís Validation Center (SNL) in Albuquerque, New

Mexico. Recent evaluation of an ATE system revealed that interpretation of the

test results could result in the identification of many wiring faults, by type and

location that were not visually observable. However, a trained system operator

was required to differentiate installation characteristics (such as clamps and

other structure adjacent to wires) from defects. The FAA concluded that further

research should be initiated to optimize the test systems for commercial

aircraft use and to develop appropriate acceptance and failure criteria for the

electrical properties of a wire. 

ï The U.S. Navy contracted with Lectromec to compare visual wire inspections

with an electronic wire inspection technique. According to the resultant report

(Report N193-RPT14MY9), during the inspections of U.S. Navy P-3 airplanes,

visual inspections detected only 25 to 39 percent of the defects that were

subsequently identified using the electronic inspection technique. 

ï The U.S. Navy has been developing an aircraft wiring system with embedded

diagnostic and prognostic capabilities (on board ATE), which is known as

smart wire. 

ï The U.S. Navy has initiated the Aircraft Wiring and Inert Gas Generator

(AWIGG) Group and has hosted a series of meetings in which Government and

industry representatives discuss AWIGG-related problems and solutions.

Many ATE system manufacturers have provided presentations regarding ATE

products and developments. 

ï The USAF AFRL has been conducting research into ATE systems and has

issued Small Business Innovation Research contracts to fund ATE research and

development.391 The FAA and the USAFís Office of Productivity, Reliability,

Availability, and Maintainability are jointly sponsoring an effort to enhance

ATE inspection systems currently being developed. 

Additionally, in June 2000, the White House Office of Science and Technology

organized a Wire System Safety Interagency Working Group (WSSIWG) to examine the

condition of wires (and other interconnections) in various transportation and industry

modes392 and to evaluate wire inspection/maintenance methods that may help maintain

safe conditions. The WSSIWG provides an opportunity for discussion and sharing of

390 For additional information regarding the ATSRAC, see section 1.18.3.4.4.

391 According to an AFRL representative, studies of USAF wire failures and inspection techniques

indicated that maintenance/inspection personnel did not always use ATE inspection methods because of

their limitations. As a result, the USAF has been seeking more user-friendly ATE inspection techniques. 
Factual Information 196 Aircraft Accident Report

information among representatives from agencies with a wide range of possibly unrelated

uses for wire. 

1.18.3 Safety Recommendations and Responses and Actions 

Resulting From the TWA Flight 800 Accident 

1.18.3.1 December 13, 1996, Safety Recommendations Regarding Fuel Tank 

Flammability (Safety Recommendations A-96-174 through -177)

During its investigation of the TWA flight 800 accident, the Safety Board became

concerned about the operation of transport-category airplanes with potentially explosive

fuel/air mixtures in the fuel tanks, noting that this was inconsistent with the basic tenet of

transport aircraft design that no single-point failure should prevent continued safe flight.

As a result of its concerns, the Safety Board issued Safety Recommendations A-96-174

through -177 to the FAA on December 13, 1996.

Safety Recommendations A-96-174 and -175393 asked the FAA to

Require the development and implementation of design or operational changes

that will preclude the operation of transport-category airplanes with explosive

fuel/air mixtures in the fuel tanks: 

(a) Significant consideration should be given to the development of airplane

design modifications, such as nitrogen-inerting systems[394] and the addition of

insulation between heat-generating equipment and fuel tanks. Appropriate

modifications should apply to newly certificated airplanes and, where feasible, to

existing airplanes. (A-96-174)

(b) Pending implementation of design modifications, require modifications in

operational procedures to reduce the potential for explosive fuel/air mixtures in

the fuel tanks of transport-category aircraft. In the 747, consideration should be

given to refueling theÖCWT before flight whenever possible from cooler ground

fuel tanks, proper monitoring and management of the CWT fuel temperature, and

maintaining an appropriate minimum fuel quantity in the CWT. (A-96-175) 

In its February 18, 1997, response, the FAA stated that these safety

recommendations proposed major changes in the requirements for fuel tank design and

fuel management on transport-category airplanes. The FAA noted that (as previously

discussed) existing airworthiness standards assume that fuel vapor in the tank is

flammable and that design requirements dictate eliminating potential fuel tank ignition

sources. The FAA stated that controlling the flammability characteristics of fuel vapor in

airplane fuel tanks would require a major change in design concept and that to control fuel

temperatures to the extent necessary to ensure that the fuel tank vapor is nonflammable

392 For example, the WSSIWG includes representatives from the FAA, U.S. Navy, USAF, NRC, and

Department of Commerce (National Institute of Standards and Technology). 

393 Safety Recommendation A-96-175 was designated ìUrgent.î

394 Fuel tank inerting involves using inert gases in fuel tanks to lower the potential for explosions. 
Factual Information 197 Aircraft Accident Report

under all operating conditions would be a major task with significant economic

implications. The FAA indicated that to evaluate these safety recommendations

adequately, it would solicit public comments regarding the technological practicality and

economic reasonableness of the Safety Boardís recommendations. A notice to solicit

public comment was published on April 3, 1997, in 62 Federal Register (FR) 16013.

Comments were required to be submitted to the FAA by August 1, 1997.395

In a June 27, 1997, letter, the FAA stated that it shared the Safety Boardís goal of

minimizing the risk of fuel tank explosions and believed that reducing fuel volatility and

ignition sources was the most promising path to this objective. However, the FAA

indicated that more research on the chemistry of fuel vapor and the energy needed to ignite

the vapor and additional flight tests to obtain data on fuel and vapor temperatures were

needed to determine the best actions to be taken. The FAA also stated that it was pursuing

an aggressive program to further minimize ignition sources within fuel tanks, including

providing ground-fault protection to wires in fuel tanks and more focused maintenance

programs for fuel tank components. 

On July 1, 1997, the Safety Board advised the FAA that Safety Recommendation

A-96-174 addressed the current FAA policy of preventing fuel tank explosions by

eliminating ignition sources because this policy has not been 100 percent effective. The

Boardís letter further noted that it is unlikely that TWA flight 800 would have experienced

the catastrophic explosion and breakup if the CWT had not contained explosive fuel/air

vapor. The Board indicated its belief that it is unacceptable to allow explosive fuel/air

vapor to exist in fuel tanks of passenger-carrying aircraft. The Board stated that, although

it was concerned about the length of time it took for the FAA to publish the request for

comments, it recognized the value in the FAA seeking and evaluating input from the

industry before mandating policy change that could require significant modifications to

aircraft designs. Therefore, Safety Recommendation A-96-174 was classified ìOpenó

Acceptable Response.î

Regarding Safety Recommendation A-96-175, the Safety Board noted that it

called for operational changes that could be instituted quickly and that neither limiting air

conditioning use on the ground nor reapportioning initial fuel loading to the CWT from

the wing tanks would require the purchase of additional equipment or fuel. The Board

395 According to this notice to solicit public comment, the FAA had initiated research into the feasibility

of nitrogen inerting as a method of ìreducing the likelihood of fuel tank explosion due to post crash ground

fireî in 1969. This research involved the installation and use of a nitrogen inerting system in a DC-9. The

FAA concluded that although ìthe system provided adequate inerting of the fuel tanksÖ.the costs of such a

system were shown to outweigh the benefits at that time.î The notice to solicit public comment also cited

Safety Recommendation A-71-59, which was issued by the Safety Board in 1971 as a result of an accident in

New Haven, Connecticut, in which 27 of 28 airplane occupants survived the initial ground impact but died

as a result of the postcrash fire/explosion. Safety Recommendation A-71-59 urged the FAA to ìrequire fuel

system fire safety devices which will be effective in prevention and control of both in flight and postcrash

fuel system fires and explosions.î On November 13, 1985, Safety Recommendation A-71-59 was classified

ìClosedóAcceptable Action.î 

The notice to solicit public comment also stated that the military has used nitrogen inerting systems in

C-5 and C-17 military transport-category airplanes, the F-22 fighter, and the V-22 tiltrotor. The military also

uses foam filler explosion suppression systems in several of its airplanes. 
Factual Information 198 Aircraft Accident Report

noted that, although reducing fuel tank temperatures by only a few degrees may seem

insignificant, incremental reductions of fuel/air temperatures could keep them below the

lower explosive limits of the fuel. Because the FAA had not initiated any immediate action

in response to Safety Recommendation A-96-175, it was classified ìOpenóUnacceptable

Response.î 

In a December 3, 1997, letter, the FAA stated that it had received 976 pages of

comments from 340 commenters regarding the Safety Boardís recommendations in

response to its notice to solicit public comment. The FAA indicated that comments from

the public generally expressed a concern that action be taken to improve safety, whereas

industry comments focused on the technical merits and costs associated with

implementing the recommendations. The FAA noted that commenters did not agree on the

reliability of possible long-term methods of reducing or eliminating explosive fuel/air

mixtures or on the weight and cost implications. The FAA indicated that it intended to

establish an ARAC working group396 to evaluate both reducing or eliminating explosive

fuel/air vapors and potential ignition sources within airplane fuel tanks and to identify

specific methods to implement improvements in transport airplane fuel tank safety for the

current and future fleet. The FAA stated that the ARAC working group would be given 6

months to provide the FAA with a report describing all technical issues and specific

solutions. 

Regarding short-term methods of reducing exposure time to explosive fuel/air

mixtures, the FAA stated that its analysis indicated that loading cool fuel into the CWT

and controlling the use of the air conditioning packs on the ground would have little effect

on the overall level of exposure to operation with explosive fuel/air mixtures and that it

did not see a significant safety benefit from adding fuel to the CWT when it would

normally be empty. However, the FAA stated that removing heat from the pack bay

through dedicated ventilation, coupled with insulating the tank from heat sources, may

significantly reduce the exposure to operation with explosive fuel/air mixtures. The FAA

indicated that design modifications, such as insulation or inerting, were among the

long-term solutions that would be evaluated by the ARAC working group.

The FAA also stated that, in the short-term, it would further reduce the likelihood

of fuel tank explosions by the following actions:

ï proposing to require periodic inspections of 747 fuel tanks and equipment

inside fuel tanks;

ï continuing to correct design or maintenance-related deficiencies in 747 fuel

tanks that may lead to ignition sources being present; and 

396 The FAA stated that the ARAC working group would be composed of industry experts, as well as a

number of passenger, union, and public interest group members. 
Factual Information 199 Aircraft Accident Report

ï proposing to require, for the fleet of all large air transport airplanes, that

(1) each type certificate holder develop a fuel tank maintenance and inspection

program, (2) each operator have an FAA-approved fuel system maintenance

program, (3) original certification compliance findings be reviewed to

revalidate that failures within the fuel system will not result in ignition sources,

and (4) operation of electrically driven fuel pumps in fuel tanks with adjacent

heat sources be prohibited when the pump inlet is not fully submerged in liquid

fuel, unless a flame arrestor is installed in the fuel pump inlet line.

In a March 3, 1998, letter, the FAA indicated that the notice to establish the new

ARAC working group (the FTHWG) was published in the FR on January 23, 1998. The

notice gave the FTHWG until July 23, 1998, to provide the FAA with a technical report.397

The FAA further noted that Safety Board-sponsored testing at CIT indicated that reducing

the temperature of the fuel in the CWT significantly increased the energy level required to

ignite the fuel and that this information conflicted with industry-standard data traditionally

used by the military and the FAA in establishing the flammability of vapors within fuel

tanks. The FAA stated that it had established a team of fuel experts to review all available

data on the relationship of fuel temperature to MIE and that a report of the teamís

conclusions and recommendations was expected by the end of March 1998. The FAA

further stated that, upon review of that report, it intended to reevaluate whether a

significant safety improvement could be achieved through practical changes in fuel

loading and fuel usage procedures. 

The FAA also reported on the status of the short-term actions (outlined it its

December 3, 1997, letter) to reduce the likelihood of fuel tank explosions. Specifically, in

addition to several planned actions, the FAA stated that it had already completed the

following actions: 

ï Issued AD 97-26-07, effective December 29, 1997 (superseding

AD 96-26-06), which required a repetitive inspection of Teflon sleeves that

protect wiring to the outboard main tank boost pumps on all 747 series

airplanes. (The FAA noted that similar AD action was planned for 767 series

airplanes.) 

ï Issued NPRM 97-NM-272-AD on November 26, 1997, which proposed to

require modification of 747 FQIS wiring to incorporate separation, shielding,

and/or electrical transient suppression features to prevent electrical signals

with excessive energy from entering the fuel tanks. (The FAA noted that a

similar NPRM was being developed for FQIS wiring on 737s.)

In a September 21, 1999, response, the Safety Board expressed concern that in the

2 1/2 years since the safety recommendations were issued, no design modifications to

airplanes had been made. Noting that the FAA had taken no action toward implementing

Safety Recommendation A-96-174, and it appeared that no action was contemplated, the

397 For additional information regarding the FTHWGís technical report, conclusions, and

recommendations, see section 1.18.3.4.2. 
Factual Information 200 Aircraft Accident Report

Board classified the recommendation ìOpenóUnacceptable Response.î With regard to

short-term actions, the Board noted that its July 1997 flight tests had demonstrated that the

potential flammability of vapors in CWTs could be reduced through fuel management and

limiting the use of air conditioning packs. The Board stated that, although the overall

reduction in flammability was not as great as originally thought, the reductions

represented a safety improvement that could be achieved at minimal cost and in a very

short time. Because the FAA had not taken any action directly responding to Safety

Recommendation A-96-175 in nearly 3 years, the Board classified it ìOpenó

Unacceptable Response.î

In a November 3, 1999, letter, the FAA noted that the FTHWG, whose report had

been issued on July 23, 1998, concluded that reducing flammability in all fuel tanks to the

level of wing tanks on most airplanes (which the FTHWG calculated to be approximately

7 percent of the fleetwide operational time) as a worthwhile goal. After evaluating various

design concepts, the FTHWG concluded that retrofitting the existing fleet was not

currently feasible but that changes could be made for new airplane designs that would

reduce the flammability of CWTs. The FAA stated that it may not agree with everything

contained in the FTHWGís report398 but that it did agree with the conclusion that further

study is needed to determine if a method can be shown to be feasible that would eliminate

or reduce the flammability of fuel tanks on in-service airplanes. The FAA indicated that it

had therefore initiated studies of directed ventilation and ground-based inerting methods

and that, after equipment requirements for ground-based inerting had been determined, it

intended to task the ARAC to perform a detailed study of inerting options.

The FAA also stated that, on the basis of conclusions of a team of fuel experts and

the FTHWG, it had determined that the information available still showed that there was

no significant safety benefit from adding fuel to 747 CWTs. However, the FAA indicated

that it was continuing to take short-term actions to reduce the likelihood of fuel tank

explosions by proposing airworthiness actions to correct any design or

maintenance-related deficiencies in fuel tank systems that may lead to an ignition source

being present.

In a May 17, 2000, letter, the FAA indicated that it was continuing to make

progress toward taking action to significantly reduce the exposure of transport-category

airplane fuel tanks to flammable vapors in both newly certificated and existing airplanes

and that methods under evaluation were cooling of the CWT using directed ventilation

and fuel tank inerting using on-board nitrogen or ground-based nitrogen sources.

According to the FAA, initial temperature data provided by airplane manufacturers

indicated that directed ventilation systems would not reduce the temperature of heated

CWTs as much as previously expected or predicted in the FTHWGís report. However, the

FAA stated that it would reach a final conclusion after it received additional temperature

data.

398 During a February 29, 2000, meeting with Safety Board staff, FAA staff indicated that the FAA did

not intend to endorse the FTHWGís proposed exposure criteria of 7 percent and would convene another

ARAC to study methods of protecting fuel tank vapors. The tasking for this ARAC was published in the FR

on July 14, 2000. 
Factual Information 201 Aircraft Accident Report

The FAA also stated that its study indicated that the cost to implement

ground-based inerting was significantly less than the FTHWGís estimate. The FAAís

analysis indicated that ground-based inerting would reduce the fleetwide average

flammability exposure for affected CWTs to approximately 2 percent. The FAA indicated

that it planned to conduct ground and flight tests to determine how long ground-based

inerting would remain effective during flight and whether vent system modifications

would be required to ensure its effectiveness. The FAA stated that these tests and its

ground-based inerting report would be completed by mid-2000. The FAA further stated

that a new ARAC working group would be established to perform a detailed study of

ground-based inerting and on-board inerting methods for in-service, newly manufactured,

and newly designed airplanes and to develop further data required for the FAA to

determine if a method of inerting fuel tanks could be practical for these airplanes. The

FAA also stated that on October 29, 1999, it had proposed regulatory changes (in NPRM

99-18)399 that would require minimizing the flammability of fuel tanks in newly designed

transport-category airplanes. The FAA stated that the proposed regulation was intended as

a short-term measure to require that fuel tanks are not heated and that they cool to below

the ignition temperature at a rate equivalent to that of a wing tank. 

With regard to operational procedures, the FAA stated that, in response to recent

suggestions from Safety Board staff, it had evaluated the use of ground sources for

conditioned air instead of using the airplaneís air conditioning packs and had determined

that this would provide a greater reduction in the exposure of those CWTs to flammable

vapors than adding cooler fuel to the CWT before flight. The FAA further stated that its

analysis indicated that the use of ground-conditioned air on days when temperatures

exceeded 60° F400 would reduce the exposure from approximately 35 percent401 to

approximately 25 percent, whereas adding cooler fuel to the CWT would result in a

reduction to only approximately 30 percent. The FAA also determined that a combination

of refueling the CWT with cooler fuel and using ground-conditioned air did not result in

any additional reduction in exposure to flammability. The FAA noted that a May 5, 2000,

Boeing SL, dated May 5, 2000, recommended that operators use ground-conditioned air

when available and practical and when ambient ground temperatures are greater than

60° F and that they evaluate opening the pack bay doors to provide additional ventilation.

The FAA stated that it would issue an information bulletin to FAA representatives

assigned to each air carrier encouraging the application of the Boeing SL.

Safety Recommendations A-96-174 and -175 and their current classifications are

discussed further in section 2.4 of this report.

399 For additional information about NPRM 99-18, see section 1.18.3.4.3. 

400 The FAA stated that analysis showed that using ground-conditioned air would provide little or no

reduction in CWT flammability, which would already be low, when ambient temperatures are below 60° F.

401 Although the FTHWG had estimated that CWTs with air conditioning packs located beneath them

would be flammable an average of 30 percent of the fleetwide operational time, the FAAís estimate of

35 percent was based on an analysis that considered the results of research completed after the FTHWG.
Factual Information 202 Aircraft Accident Report

Safety Recommendation A-96-176 asked the FAA to 

Require that the 747 Flight Handbooks of TWA and other operators of 747s and

other aircraft in which fuel tank temperature cannot be determined by flight crews

be immediately revised to reflect the increases inÖCWT fuel temperatures found

by flight tests, including operational procedures to reduce the potential for

exceeding CWT temperature limits. 

(Although the FAAís February 18 and June 27, 1997, letters referenced Safety

Recommendations A-96-174 through -177, they did not specifically address Safety

Recommendation A-96-176.) 

In its July 1, 1997, letter to the FAA, the Safety Board noted that correcting these

and similar flight handbook errors would not have required a delay while the FAA sought

public comment. Accordingly, Safety Recommendation A-96-176 was classified ìOpenó

Unacceptable Response.î

In its December 3, 1997, letter, the FAA acknowledged that the Safety Boardís

flight tests had shown that it was possible for the temperature of the fuel in the CWTs of

Boeing 747s to rise significantly during extended periods of air conditioning pack

operation on the ground. However, the FAA stated that it had not required that such

information be provided to operators in the FAA-approved aircraft flight manual (AFM)

and that the information in the Boeing 747 TWA Flight Handbook is considered by the

FAA to be advisory only.

The FAA also noted (in its December 3, 1997, and March 3, 1998, letters) that the

fuel tank temperature limits given in the FAA-approved AFM for the Boeing 747 were

intended to prevent freezing of the fuel and to ensure that the fuel feed system will

perform in hot weather without any anomalous operation caused by excessive fuel

vapor/liquid ratios within the fuel feed system, not to preclude operation of the airplane

with a flammable fuel/air mixture in the fuel tanks. 

In a September 21, 1999, letter, the Safety Board noted that the information in the

TWA 747 Flight Handbook about temperature limits was incorrect in that it states that air

conditioning pack operation can elevate the CWT temperature by 10° to 20° F, whereas

flight tests found increases from the baseline fuel tank temperature of 40° to 60° F during

air conditioning pack operation. Because the FAA had taken no action directly responding

to Safety Recommendation A-96-176, it was again classified ìOpenóUnacceptable

Response.î

In its May 17, 2000, letter, the FAA noted that TWA no longer operates 747s and

that the TWA flight handbooks that were the subject of this recommendation are no longer

active. According to the FAA, Boeing had reviewed its flight and operations manuals and

polled operators regarding the content of their flight handbooks and determined that none

of those documents provided CWT temperature correction information for flight crews.402

On the basis of this information, in an August 3, 2000, letter, the Safety Board classified

Safety Recommendation A-96-176 ìClosedóNo Longer Applicable.î 
Factual Information 203 Aircraft Accident Report

Safety Recommendation A-96-177 asked the FAA to 

Require modification of theÖ[CWT] of 747 airplanes and the fuel tanks of other

airplanes that are located near heat sources to incorporate temperature probes and

cockpit fuel tank temperature displays to permit determination of fuel tank

temperatures. 

(Although the FAAís February 18 and June 27, 1997, letters referenced Safety

Recommendations A-96-174 through -177, they did not specifically address Safety

Recommendation A-96-177.) 

In its July 1, 1997, letter to the FAA, the Safety Board noted that without

temperature probes and indicators, 747 flight crews could not be alerted to hazardous

temperatures in the CWT. The Board stated that these changes, if implemented, would

allow flight crews to reduce exposure to high CWT temperatures and that neither change

would have significant economic impact. Although the Board expressed concern about the

length of time it took the FAA to publish its request for comments on the Boardís

recommendations, it recognized the value in seeking and evaluating input from the

industry. Accordingly, Safety Recommendation A-96-177 was classified ìOpenó

Acceptable Response.î

As previously discussed, in its December 3, 1997, letter, the FAA stated that the

purpose of the fuel tank temperature limits in Boeing's AFM was not for the management

of fuel tank temperature to preclude the presence of a flammable mixture in the fuel tank

ullage and that it considered the current tank temperature indication scheme used on

transport-category airplanes to be adequate for the purposes for which it was intended.

The FAA indicated that comments on this recommendation from industry, operators, and

manufacturers questioned the usefulness of providing such information to flight crews. On

the basis of those comments, the FAA indicated that the management of fuel tank

temperature on existing airplane configurations through the creation of a tank temperature

ìred-lineî associated with tank flammability is not considered to be practical. It also noted

that the addition of another set of electrical components and wiring inside the fuel tank

would create another possible source of tank ignition.

In a March 3, 1998, letter, the FAA stated that, as additional technical data and

information were developed through the ongoing activities (including the FTHWG and

Safety Board-sponsored fuel research at CIT), it would further evaluate the need to take

the specific actions identified by the Board in Safety Recommendation A-96-177. In its

September 21, 1999, letter, the Safety Board stated that, because the FAA had taken no

action to implement Safety Recommendation A-96-177, it was classified ìOpenó

Unacceptable Response.î

In a November 3, 1999, letter, the FAA indicated that the work of the FTHWG and

researchers at CIT was complete. The FAA stated that it had concluded that no practical

402 Boeing found that one non-U.S. operator had a note similar to the one found in the TWA Flight

Handbook, and it requested that the operator remove that note.
Factual Information 204 Aircraft Accident Report

means of reducing fuel temperatures in fuel tanks existed but that, as stated in response to

Safety Recommendation A-96-174, it had begun studies of directed ventilation methods to

determine if they could be effective in cooling CWTs and be feasible for retrofit on the

existing fleet of transport-category airplanes. The FAA stated that it would reevaluate this

recommendation after the conclusion of the studies of methods that could significantly

reduce the flammability of fuel tanks.

The FAA stated that, although certain military airplanes incorporate fuel

temperature indications with cockpit temperature displays for each tank, those displays are

not used to determine fuel tank flammability; instead, they are used to monitor fuel tank

temperatures during extended ground operation to ensure that the fuel temperatures in the

main fuel tanks (wing tanks) do not rise to a range that would cause degraded fuel pump

performance or pump cavitation when the engines are operating at high thrust levels that

require high fuel flow from the fuel pumps. The FAA reiterated its belief that no benefit

would be gained from additional fuel tank temperature probes.

In its May 17, 2000, letter, the FAA stated that its evaluations of directed

ventilation and fuel tank inerting systems indicated that flight crews would not require

CWT temperature information in connection with those methods of controlling

flammability. Further, the FAA stated that the use of ground-conditioned air would not

require that information either.

In an August 3, 2000, letter, the Safety Board agreed that using these methods to

reduce flammability could obviate flight crews needing fuel tank temperature information

but noted that neither directed ventilation nor fuel tank inerting systems are currently in

use and, although a Boeing SL recommends the use of ground-conditioned air when

available, it is not required. Accordingly, pending implementation of adequate methods

for reducing fuel tank flammability or for determining fuel tank temperatures, Safety

Recommendation A-96-176 was classified ìOpenóUnacceptable Response.î

1.18.3.2 February 18, 1997, Safety Recommendation Regarding the 

Handling and Placement of Explosive Training Aids (Safety 

Recommendation A-97-11)

During the TWA flight 800 investigation, the Safety Board became aware that on

June 10, 1996, a dog handler, working under the auspices of the FAAís Explosives

Detection Canine Team Program, had spilled trace amounts of explosives while placing

training aids on board the accident airplane during a proficiency training exercise. As a

result of this finding, on February 18, 1997, the Board issued Safety Recommendation

A-97-11, which asked the FAA to

Develop and implement procedures, including a checklist of safety-related items,

for the handling and placement of explosive training aids by K-9 explosives

detection teams to prevent contamination of aircraft and airport facilities and to

ensure an effective K-9 explosives detection program. 
Factual Information 205 Aircraft Accident Report

In its April 18, 1997, February 9 and December 14, 1998, letters, the FAA stated

that it had developed several initiatives to safeguard and account for canine explosives

training aids used by participants in the FAAís Explosives Detection Canine Team

Program. For example, the FAA indicated that on May 19, 1997, it initiated a national

database to record explosives training conducted on aircraft. In addition, during August

and September 1997 and January 1998, the FAA provided updated and detailed training

for all FAA canine coordinators regarding proper safe handling, storage, and

accountability for explosives training aids. The FAAís letter further stated that it procured

explosives magazines approved by the ATF for each of the FAAís Explosives Detection

Canine Team Program participants. According to the FAA, these magazines are secured

with two high-security locks and provide a greater degree of security while reducing

contamination problems. 

The FAA also indicated that it had developed procedures for the use of barrier

material(s) to prevent leaving explosives residues on surface areas when placing

explosives training aids on aircraft or within the airport environment. In addition, the FAA

developed an Explosives Detection Canine Team Training Log (in lieu of the checklist

suggested in Safety Recommendation A-97-11) to aid in monitoring training requirements

by requiring documentation of the type and number of training aids used and verification

that the training aids were removed from the aircraft/airport environment. Further, the

FAAís letter indicated that on October 28, 1998, FAA staff briefed the Safety Board on the

contents of the FAA Explosives Detection Canine Team Program Standard Operating

Practices and explained how the procedures described therein addressed the intent of

Safety Recommendation A-97-11. 

On February 8, 1999, the Safety Board responded that, because the FAAís actions

satisfied the intent of Safety Recommendation A-97-11, it was classified ìClosedó

Acceptable Action.î 

1.18.3.3 April 7, 1998, Safety Recommendations Regarding Fuel Quantity 

Indication System Wiring and Components (Safety Recommendations 

A-98-34 through -39)

As a result of its concerns about the potential for excessive energy to enter airplane

fuel tanks through the FQIS and result in possible fuel tank ignition sources, the Safety

Board issued Safety Recommendations A-98-34 through -39 on April 7, 1998. 

Safety Recommendation A-98-34 asked the FAA to

Issue, as soon as possible, anÖ[AD] to require a detailed inspection ofÖ [FQIS]

wiring in Boeing 747-100, -200, and -300 series airplane fuel tanks for damage

and the replacement or the repair of any wires found to be damaged. Wires on

Honeywell Series 1-3 probes and compensators should be removed for

examination. 
Factual Information 206 Aircraft Accident Report

Safety Recommendation A-98-35 asked the FAA to

Issue anÖ[AD] to require the earliest possible replacement of the Honeywell

Corporation Series 1-3 terminal blocks used on Boeing 747 fuel probes with

terminal blocks that do not have knurled surfaces or sharp edges that may

damageÖ[FQIS] wiring. 

In a June 10, 1998, letter, the FAA stated that Boeing had issued SB 747-28-2205,

Revision 1, calling for inspections and testing to verify that the wiring, tubing, and

component installations inside the CWT are in satisfactory condition and electrically

bonded to the airplane structure. The FAA indicated that Boeing had also issued SBs

calling for the replacement of Series 1-3 probes with Series 4 or later probes

(SB 747-28-A2208) and installation of a flame arrestor in the inlet line of the electrical

motor-operated scavenge pump located in the CWT (SB 747-28-A2210). The FAA stated

that it would issue an NPRM proposing to require compliance with these SBs, including a

requirement to replace CWT fuel probes at a specified period of time. 

On July 24, 1998, the FAA issued NPRM 98-NM-163-AD, which proposed to

require compliance with SBs 747-28-2205, Revision 1, and 747-28-A2208 and to require

replacement of FQIS components within 20 years of manufacture. In a February 9, 1999,

letter, the Safety Board stated that it did not believe the referenced Boeing SBs contained

sufficiently detailed instructions to ensure that the inspections were adequately thorough.

For example, the Board noted that, although much of the damage it observed during its

airplane wiring inspections was not visible until wiring was removed from its installation,

the SBs did not specify the removal of wires for inspection. The Board further noted that

the SBs did not address the issue of migration (cold-flow) of insulation material or the

presence of localized blackening on the wire surfaces, which might indicate breaches in

the insulation and, thus arcing, inside the tank. This, in turn, could ignite the flammable

fuel vapors.403 The Board also noted that the FAA had provided no statistical basis for its

proposed requirement to replace FQIS components within 20 years of their manufacture.

Therefore, pending information from the FAA indicating that adequate and thorough

inspections would be required and described and demonstrating the appropriateness of a

20-year service life for FQIS components, Safety Recommendation A-98-34 was

classified ìOpenóUnacceptable Response.î Pending final rulemaking to remove Series

1-3 terminal blocks (which have knurled surfaces) from the 747 fuel tanks within 2 years

of the ADís effective date, Safety Recommendation A-98-35 was classified ìOpenó

Acceptable Response.î 

In a November 3, 1999, letter, the FAA acknowledged the Safety Boardís concern

about hidden damage to wires going undetected during inspections but expressed concerns

that the removal and replacement process might cause additional wire damage. The FAA

stated that it believed that ìinspections that minimize the amount of disturbance of wiring

are the best compromise between the risks associated with wire removal and hidden

damage.î The FAA indicated that it believed that the in-tank inspections and

403 Conductors exposed as a result of migration or displacement of the wire insulation (cold-flow) was

one of the suspected ignition mechanisms for the 1990 Philippine Airlines 737 CWT explosion. 
Factual Information 207 Aircraft Accident Report

modifications (including the replacement of FQIS probes with knurled surfaces or sharp

edges) required by AD 99-08-02 R1,404 issued on June 29, 1999, in combination with the

shielding/separation of FQIS wiring outside the fuel tanks required by AD 98-20-40405

would provide the level of safety intended by Safety Recommendation A-98-34. The FAA

also stated that, on October 26, 1999, it issued a Special Federal Aviation Regulation

(SFAR),406 which proposed to require a design review revalidating the fuel tank system

designs to ensure that failures could not create ignition sources within the fuel tanks and to

require manufacturers to design specific fuel tank maintenance and inspection programs to

ensure the continued safety of fuel tank systems.407

Regarding the proposed 20-year life limit for FQIS components, the FAA agreed

that there was little technical basis for the selection of that life limit and stated that it had

hoped to gather additional information on an appropriate life limit through the NPRM

comment process. However, on the basis of the comments that it had received, the FAA

determined that additional study of sulfide deposits was necessary before any life limit

could be imposed.

On the basis of the FAAís actions, in an August 3, 2000, letter, the Safety Board

classified Safety Recommendations A-98-34 and -35 ìClosedóAcceptable Action.î

Safety Recommendation A-98-36 asked the FAA to

Conduct a survey ofÖ[FQIS] probes and wires in Boeing 747s equipped with

systems other than Honeywell Series 1-3 probes and compensators and in other

model airplanes that are used in Title 14Ö[CFR] Part 121 service to determine

whether potential fuel tank ignition sources exist that are similar to those found in

the Boeing 747. The survey should include removing wires from fuel probes and

examining the wires for damage. Repair or replacement procedures for any

damaged wires that are found should be developed. 

404 In response to AD 99-08-02 R1, on June 27, 1997, Boeing issued SB 747-28-2205, which described

fuel tank inspection procedures for 747 operators. Eight of the nine inspection tasks described involved

visual inspections, and none of the inspection tasks required removal of wiring from fuel quantity probes or

compensators. (During this investigation, Safety Board investigators repeatedly found that it was not

possible to detect damage on the concealed side of wiring.) On September 25, 1997, Boeing issued a Notice

of Status Change for SB 747-28-2205. And on October 27, 1997, Boeing notified 747 operators by telex

(M-7220-970-1725) that it was developing a new SB that would provide operators with details regarding the

inspection of 747 fuel quantity probes and FQIS wires. In an October 30, 1997, letter to the Board, Boeing

indicated that the new SB would require the removal of wiring from terminal blocks for inspection,

replacement of Series 1-3 fuel quantity probes, reporting of damage found at Series 4 and subsequent fuel

quantity probes, replacement of certain CWT FQIS wire harnesses, inspection for proper wire routing and

existing damage, establishment of an electrical resistance check of very low voltage, and establishment of

standards for permissible FQIS repairs.

405 AD 98-20-40 is discussed further in connection with Safety Recommendation A-98-98.

406 For more information about the proposed SFAR, which was issued as part of NPRM 99-18, see

section 1.18.3.4.3. 

407 The FAA reiterated these actions in its May 17, 2000, letter.
Factual Information 208 Aircraft Accident Report

In its June 10, 1998, letter, the FAA indicated that it was working with an industry

team to develop plans to survey the condition of fuel tank systems on large

transport-category airplanes and that the survey would include a review of historical data,

including maintenance data. The FAA stated that the industry team had indicated that

although it did not intend to remove wiring from FQIS probes (for fear of causing damage

that would go undetected), FQIS wires in fuel tanks would be inspected for damage from

chafing and abrasion under the clamps. The FAA indicated that it would review the

industry teamís plans for the survey, evaluate the benefit of removing wires from fuel

probes, and monitor the results of the inspections. The FAA further indicated that it

planned to issue an NPRM proposing an SFAR that would require a design review of the

fuel tank system in all large transport-category airplane fleets. 

In its February 9, 1999, response, the Safety Board expressed its concern that the

industry team did not intend to remove the FQIS wires to avoid the potential for

undetected damage to the wiring. The Board emphasized that in several cases its staff

detected damage to 747 wire insulation only after the wires were removed from their

installed locations; the damage subsequently observed included compression ruptures,

cold-flow, chafes, and cuts. The Board urged the FAA to reconsider the benefits of

removing the wiring from the fuel probes to allow a thorough inspection for damage.

Pending the results of the FAAís inspections and the issuance of the SFAR, Safety

Recommendation A-98-36 was classified ìOpenóAcceptable Response.î 

In a November 3, 1999, response to the Safety Board, the FAA indicated that

although the manufacturersí instructions for the surveys of FQIS wiring in

transport-category airplanes did not include removal of FQIS wiring, all FQIS wires in

fuel tanks would be visually inspected for damage from chafing, abrasion under clamps,

etc. The FAA further stated that it did not see a need to mandate removal of FQIS wires in

airplanes other than 747s with FQIS probes that have knurled surfaces or sharp edges. The

FAA indicated that it would continue to monitor the industry teamís findings and evaluate

the need for future actions.

In November 3, 1999, and May 17, 2000, letters, the FAA noted that on October

26, 1999, it had issued the planned SFAR, which proposed a design review to revalidate

the fuel tank system designs to ensure that failures could not create ignition sources within

the fuel tanks.408

In an August 3, 2000, letter, the Safety Board stated that, pending completion of

the inspections currently in progress and the proposed fault tolerance evaluation and the

implementation of the maintenance programs that will result from the SFAR, Safety

Recommendation A-98-36 was classified ìOpenóAcceptable Response.î

408 For further details about the SFAR, see section 1.18.3.4.3.
Factual Information 209 Aircraft Accident Report

Safety Recommendation A-98-37 asked the FAA to

Require research into copper-sulfide deposits onÖ[FQIS] parts in fuel tanks to

determine the levels of deposits that may be hazardous, how to inspect and clean

the deposits, and when to replace the components. 

In its June 10, 1998, response, the FAA indicated that it was establishing a fuel

tank system copper-sulfide research program, which it intended to conduct in partnership

with the DoD. The FAA stated that the research program would address the identification

of the conditions required for film deposit formation, determination of film properties, and

diagnostics and film deposit removal. 

In February 9, 1999, letter, the Safety Board stated that it supported the FAAís plan

and provided the FAA with information gathered during the TWA flight 800 investigation

that might be useful to its research. Pending further information about the FAAís current

and planned actions, Safety Recommendation A-98-37 was classified ìOpenóAcceptable

Response.î 

In a November 3, 1999, letter, the FAA indicated that in April 1999, it had initiated

a research program (through the FAA Aviation Assurance Center of Excellence Program)

into the effects of sulfur-compound deposits on fuel system components. Participants in

the research program included the FAA, Boeing, BFGoodrich, Arizona State University,

and the UDRI. The Safety Board attended meetings as an observer and provided

information pertaining to sulfides that was obtained during the TWA flight 800

investigation. The FAA indicated that it expected a report from the research team by the

end of January 2000. 

In its May 17, 2000, letter, the FAA stated that it was continuing to make progress

in understanding copper-sulfur and silver-sulfur deposits and the potential hazards that

they may cause. The FAA indicated that the researchers had grown deposits in a

laboratory but had not found indications of significant deposits occurring in in-service

airplanes to compare with the laboratory-grown deposits. The FAA indicated that the

research program had been extended so that the team could complete its evaluation.

Safety Recommendation A-98-37 and its current classification are further

discussed in section 2.3.2.2.3.1 of this report.

Safety Recommendation A-98-38 asked the FAA to

Require in Boeing 747 airplanes, and in other airplanes withÖ[FQIS] wire

installations that are corouted with wires that may be powered, the physical

separation and electrical shielding of FQIS wires to the maximum extent possible. 
Factual Information 210 Aircraft Accident Report

Safety Recommendation A-98-39 asked the FAA to

Require, in all applicable transport airplane fuel tanks,Ö[STPs] to prevent

electrical power surges from entering fuel tanks throughÖ[FQIS] wires. 

In its June 10, 1998, letter, the FAA stated that NPRM 97-NM-272, issued on

November 26, 1997, proposed to adopt an AD applicable to 747-100, -200, -300, -SP, and

-SR airplanes which would require installation of electrical transient suppression devices

(TSD) and/or the installation of shielding and separation of the electrical wiring of the

FQIS located outside the fuel tanks. The FAA stated that it also issued a similar NPRM

(98-NM-50-AD) applicable to 737-100, -200, -300, -400, and -500 series airplanes.

Regarding other transport-category airplanes, the FAA indicated that the team developed

to work on the fuel tank design review SFAR was developing advisory material for

conducting the design reviews, which would include information on the use of wire

separation and electrical shielding and/or TSDs to protect the FQIS wiring inside fuel

tanks when applicable to the different airplane designs. 

In its February 9, 1999, response, the Safety Board noted that on September 30,

1998, the FAA issued AD 98-20-40, which required the installation of shielding and

separation of the electrical wiring of the FQIS wiring located outside the fuel tanks in 747

airplanes. In addition, the Board acknowledged the FAAís NPRM proposing a similar AD

applicable to 737s. However, the Board expressed concern that the degree of separation

required by AD 98-20-40 might not adequately protect fuel tanks from FQIS wiring

deficiencies that might allow excessive energy into the tanks. The Board stated that

information from several investigations indicated that Boeingís minimum separation

standard of 1/4409 inch might be insufficient. For example, as previously discussed, when

the Board assisted in the AAIBís investigation of a January 9, 1998, wiring fire on board a

United Airlines 767, investigators found insulation melted away from wires located more

than 1 inch from the location of the electrical arc. Balls of copper were found embedded

into insulation and other materials. The Board expressed a similar concern about adequate

separation of FQIS wiring in other large transport-category airplanes, including the Airbus

300 and 310 airplanes, in which powered cables and FQIS wiring are also corouted. 

The Safety Board also noted that, although the NPRM preceding AD 98-20-40 had

proposed a requirement for installation of electrical TSDs, the AD did not require such

devices. The Board stated that it did not believe that wire separation and shielding alone

would protect against the entry of power from sources attached to FQIS wires. The Board

noted that, for example, in a 747, 115-volt a.c. power and FQIS circuits are contained in

common electrical connectors at the fuel gauges and in other components, such as the

VSO. The Board stated that surge suppression systems would protect against damage that

could affect numerous wires at once and against unforeseen circumstances. Pending

receipt of further details regarding the wiring separation that would be accomplished by

AD 98-20-40 and planned activities for other airplane models, Safety Recommendation

409 The Safety Board cited the lesser of Boeing's required separation distancesó1/4 inch in pressurized

areas and 1/2 inch in unpressurized areas.
Factual Information 211 Aircraft Accident Report

A-98-38 was classified ìOpenóAcceptable Response.î However, noting that the FAA

had apparently decided against TSDs in AD-98-20-40, Safety Recommendation A-98-39

was classified ìOpenóUnacceptable Response.î

In its November 3, 1999, and May 17, 2000, letters to the Safety Board, the FAA

acknowledged that, although the NPRM that led to AD 98-20-40 proposed TSDs as an

optional or additional requirement to the shielding and separation requirement, that

requirement was not included in the AD. The FAA stated that, when the AD was issued, it

had determined that TSD technology was not fully developed for application to in-service

airplanes and that wire separation and shielding could provide the required level of

protection. However, the FAA indicated that it had since determined that either separation

and shielding or TSD technology, executed properly, will adequately ensure that highenergy signals are not introduced on wiring inside fuel tanks. The FAA further stated that,

to comply with the ADs,410 the design must include separation and shielding of wires

between FQIS components and separation of the connectors to the components and

circuits inside the components.

The FAA stated that it was also evaluating the acceptability of TSD technology

and, although the evaluations of proposed designs were not complete, it indicated that

TSDs could be used to meet the intent of the ADs. However, the requirements for approval

of a design using TSD technology would include demonstrating that the device would

limit the transfer of electrical energy and power through the wires that enter a fuel tank to

intrinsically safe levels (that is, limit to less than spark energy and filament heating energy

limits that could create an ignition spark or a hot surface ignition source in a piece of wire,

such as a strand of steel wool, which could short between FQIS circuits in the tanks).411

For other transport-category airplanes with FQIS, the FAA stated that draft

AC 25.981.1X, prepared as guidance material for compliance with the fuel tank design

SFAR, included information on the design and maintenance lessons learned in connection

with the TWA flight 800 investigation and described the use of wire separation and

shielding, or TSD technology, to protect FQIS wiring from short-circuits to other system

wiring and other failures and malfunctions that could induce high electrical energy on

FQIS wires that enter fuel tanks.

Safety Recommendations A-98-38 and -39 and their current classifications are

further discussed in sections 2.3.2.1.2.2.3 and 2.3.2.1.2.2.4 of this report, respectively.

410 The FAA also noted that a similar AD, requiring separation and shielding of FQIS wires and circuits

on 737s, was issued on January 26, 1999.

411 For additional information, see Federal Aviation Administration. 2000. FQIS Transient Suppression

Device/Unit for Installation on Boeing Model 747-100, -200, -300, SP, and SR Series Airplane. Issue

Paper P-1. Washington, DC. This issue paper stated that the FAA would require ìno single failure, or

combination of failures not shown to be extremely improbable [not anticipated to occur during the entire

operational life of all airplanes of that type], will allow 200 microjoules and/or 30 milliamps, or greater, into

the fuel tank assuming the conditions for spark energy or filament heating exist.î 
Factual Information 212 Aircraft Accident Report

1.18.3.4 Actions Resulting From the TWA Flight 800 Accident

1.18.3.4.1 Airworthiness Directives 

As part of the TWA flight 800 investigation, the FAA examined earlier ADs for

issues that might potentially have been related to the accident. This examination revealed

that (as previously discussed) on March 8, 1979, the FAA issued AD 79-05-04, which

required (among other actions) inspections of wires in main fuel tanks on a limited

number of 747 airplanes, with installation of Teflon sleeves on the effected wires (in

accordance with Boeing ASB 747-28A2091) required to terminate repeat inspections.

AD 79-05-04 was issued as a result of the FAAís investigation of a report that fuel pump

wires in an aluminum conduit that was routed through the fuel tank from a wing spar to a

fuel pump had chafed through the insulation, resulting in electrical arcing. The arcing

from the chafed wire to the conduit created a hole; however, the arcing did not result in

fire or explosion because of the level of the fuel. After the FAA issued AD 79-05-04, an

FAA fleetwide survey indicated that wires in these conduits were chafing against the

conduit and that up to 80 percent of the total insulation thickness was missing on

numerous wires in many airplanes. On the basis of these findings, on March 19, 1979, the

FAA issued AD 79-06-02, which made the inspections and possible replacement actions

applicable to other airplanes. 

During the investigation of the TWA flight 800 accident, the FAA determined that

an environment conducive to vibration existed in the conduit and wire bundles of the fuel

boost pumps and the auxiliary fuel tank jettison pumps, which could cause abrasion of the

Teflon sleeving and subsequent abrasion of the wires in the bundles. Thus, on January 21,

1997, the FAA issued AD 96-26-06, which required the removal of wires from conduits in

all 747sí Nos. 1 and 4 main fuel tanks and inspections and replacement of damaged

sleeving or wires. The FAA superseded AD 96-26-06 with AD 97-26-07 (applicable to all

747 airplanes) on December 29, 1997. AD 97-26-07 required the inspection of the

auxiliary fuel tank jettison pumps as well as of the wiring and Teflon sleeves in conduit for

the Nos. 1 and 4 main fuel tanks. 

After these ADs were issued, the FAA became aware of ìsevere wearî of the fuel

boost pump wiring in two model 737-200 airplanes that had accumulated 54,000 and

67,000 total flight hours, respectively. Additionally, the FAA received a report about

pin-sized holes in the conduit of a 737-200 that had 75,000 total flight hours. As a result of

these reports, on April 24, 1998, Boeing issued ASB 737-28A1120, which described

procedures for removal of the 737 fuel boost pump wiring and inspection or replacement

(with the addition of Teflon sleeves) of that wiring. Further, the FAA issued Telegraphic

AD T98-10-51, which required removal of the 737 fuel boost pump wiring and inspection

or replacement (with the addition of Teflon sleeves) of that wiring and was applicable to

737-100, -200, -300, -400, and -500 airplanes with more than 50,000 flight hours. 

After it received additional reports and information about arcing on boost pump

wiring on airplanes of various ages and flight-hour accumulation, the FAA issued the

following ADs to revise the pertinent requirements: (1) AD T98-11-51, which broadened

the inspection requirements outlined in AD T98-10-51 to include wires in 737 wing
Factual Information 213 Aircraft Accident Report

conduits, on May 10, 1998; (2) AD 98-11-52, which revised the flight-hour requirement

outlined in AD T98-11-51 to include airplanes with more than 40,000 flight hours, on

June 24, 1998; and (3) AD 99-21-15, which further revised the flight-hour requirement

outlined in AD T98-11-51 to include airplanes with more than 20,000 flight hours, on

October 8, 1999. Figure 39 shows the 737 fuel pump conduit wear rate that was identified

by the FAA during representative 500 wire inspections.412 Similar data for 747 inspection

results were not released.

The FAA also received a report of corrosion on bonding jumper wires located

within a concentrated area in the wing fuel tanks on an in-service Airbus A300. Further

inspections of other Airbus model airplanes did not reveal similar extensive corrosion or

damage to bonding jumper wires; however, they did reveal evidence of the accumulation

of deposits composed of copper, sulfur, and silver (also known as silver-sulfide deposits)

around the outer braid of some jumper wires. Tests conducted by the manufacturer showed

that these deposits did not affect the bonding function of the leads. Airbus has developed a

one-time inspection SB for its airplanes to determine the extent of the silver-sulfide

deposits and to ensure that the level of jumper wire damage found on the A300 is not

widespread. 

On November 26, 1997, the FAA issued NPRM 97-NM-272-AD, which proposed

a requirement for operators of 747-100, -200, and -300 airplanes to install components for

the suppression of electrical transients and/or the installation of shielding and separation

of FQIS wiring from other airplane system wiring. On September 23, 1998, the FAA

issued AD 98-20-40, which required the installation of shielding and separation of the

FQIS wiring on 747-100, -200, and -300 airplanes. This action was intended to preclude

high levels of electrical energy from entering the airplane fuel tank wiring because of EMI

or electrical short circuits. (On April 14, 1998, the FAA issued NPRM 98-NM-50-AD,

which proposed a similar requirement for 737-100, -200, -300, -400, and -500 airplanes;

the resultant AD 99-03-04 was issued on January 26, 1999.) According to the FAA, the

FQIS wiring in all later model 737s and 747s has wire separation and fault isolation

features that may meet the intent of these ADs. Rulemaking currently under consideration

(Docket No. 99-18, discussed later in this section) will require evaluation of the more

recent designs.

On March 30, 1998, the FAA received reports of three instances of electrical

arcing within fuel pumps installed in Lockheed L-1011 fuel tanks. In one case, the

electrical arc had resulted (as previously mentioned) in a hole in the side of the pump.

According to the FAAís AD-related text, ìthe internal electrical failures in the wing fuel

tank fuel boost pumps could result in either electrical arcing or localized overheating,

which could breach the protective housing and expose it to fuel vapors and consequent

potential fire or explosion in the wing fuel tank.î On April 3, 1998, the FAA issued

emergency AD 98-08-09, which required that fuel be carried in the fuel tanks to cover the

fuel pumps thereby precluding ignition of vapors within the fuel tank (by this mechanism)

until a corrective action was developed. 

412 For additional information, see the FAAís Web site at <http://www.faa.gov/apa/737iu.htm>. 
W ear Rate Found by 737 Fuel Pum p

Conduit W ire Inspections

05

101520253035

30-40,000 40-50,000 50-60,000 60-70,000 > 7

Airplane Flight Hours 

Percentage of Wear 

Through Insulation

Upper Line Denotes Exposed Conductor, 

Lower Line Denotes 50% Insulation W ear

Figure 39. A chart showing the 737 fuel pump conduit wear rate that was identified by the FAA during represenwire inspections.
Factual Information 215 Aircraft Accident Report

The Direction Generale de LíAviation Civile (DGAC) received three reports of

CWT pumps in Airbus A300s with cracked canister legs caused by fatigue. In one case,

this led to separation of the pump canister from its lower housing. As a result of these

events, on November 30, 1998, the DGAC issued French Telegraphic

AD T98-476-272(B), which required inspection and possible replacement of the fuel

pumps on Airbus A300-B4-600R and A300-F4-600R airplanes. T98-476-272(B) stated

that the cracked canister legs could result in loss of flame trap capability and could

provide an ignition source in the CWT. As a result of the DGAC action, the FAA issued

AD T98-25-93 on December 7, 1998, requiring similar action. 

The FAA became aware of another unanticipated fuel tank system failure

condition as a result of reported incidents involving 767 airplanes in which parts from fuel

pump assemblies contacted the rotating fuel pump impeller. In two reported cases, objects

from a broken fuel pump inlet diffuser assembly were ingested into the fuel pump, causing

damage to the pump impeller and pump housing. Because this condition could have

resulted in sparks or hot debris in the presence of fuel vapor, the FAA issued

AD 97-19-15, which required revision of the 767 AFM to include procedures to switch off

the fuel pumps when the CWT is nearly empty. According to the FAA, the intent of this

action is to maintain liquid fuel over the pump inlet so that any debris generated by a

failed fuel pump would not come into contact with fuel vapors. 

1.18.3.4.2 Aviation Rulemaking Advisory Committee/Fuel Tank 

Harmonization Working Group 

On April 3, 1997, the FAA published a notice in the FR that requested comments

about the Safety Boardís Safety Recommendations A-96-174 through -177 regarding

reduced flammability. As previously discussed, on the basis of its review of comments

that it had received, the FAA determined that additional information was needed before it

would initiate rulemaking action to address the Boardís recommendations. Therefore, on

January 23, 1998, the FAA established an ARAC working group, the FTHWG, which was

tasked with gathering the information needed. Representatives from the Aerospace

Industries Association, Air Transport Association of America (ATA), International Air

Transport Association, European Association of Aerospace Industries, Air Line Pilots

Association, API, General Aviation Manufacturers Association, Joint Aviation Authorities

(JAA), Aviation Consumer Action Project (ACAP), and FAA were involved in the

FTHWG. 

The FTHWG was tasked with recommending regulatory changes that would

ìeliminate or significantly reduce the hazards associated with explosive vapors in

transport-category airplane fuel tanks [of]Önew type designs, in-production airplanes and

the existing fleet.î Its objectives included the following: (1) analyze the history of the

world transport-category aircraft fleet, the safety status of the existing fleet, various means

of reducing exposure to flammable fuel vapors, and the means to eliminate the resultant

hazard if ignition does occur; (2) recommend regulatory text for new rulemaking aimed at

controlling flammability of fuel vapors in fuel tanks; and (3) assess the cost benefit of

those means. 
Factual Information 216 Aircraft Accident Report

On July 23, 1998, the FTHWG submitted its final report to the FAA. The report

indicated that the FTHWGís studies showed that flammability exposure varied across

airplane types and depended on fuel tank location. According to the report, some fuel

tanks already have a low exposure to flammable conditions. For example, the FTHWG

calculated that wing tanks are flammable approximately 7 percent of the fleetwide

operational time. Further, results from the FTHWGís thermal analysis indicated that the

flammability exposure for fuel tanks that were not adjacent to such heat sources

approximated that of wing tanks, whereas CWTs heated by air conditioning equipment

located beneath them are flammable an average of 30 percent of the fleetwide operating

time.413 The FTHWG concluded that the safety record of wing tanks was adequate and if

the same level could be achieved in all CWTs, an adequate overall safety objective would

be achieved. Accordingly, reducing flammability in all fuel tanks to the level of the wing

tanks on most airplanes was seen by the FTHWG as a worthwhile goal. Various possible

means to achieve this goal were evaluated for technical and economic merits, and the

FTHWG reached the following conclusions: 

ï Techniques to reduce or eliminate heat input to the tanks from nearby heat

sources were evaluated. Of these techniques, directed ventilation and

relocation of the significant heat sources reduce the exposure to an acceptable

level. However, relocation is only feasible for new airplane designs. Directed

ventilation for in-service aircraft is estimated to have an overall cost of $3.5

billion over a 10-year period. 

ï To reach the goal by changing fuel properties, a minimum flash point

specification of 140° F would be required. A change of this magnitude falls

outside of the current experience base and may require engine re-design/requalification. The overall fuel manufacturing cost increase for a 10-year period

is estimated at $15 billion in the USA and $60 billion for the rest of the world

and could result in a significant shortfall of jet fuel.

ï Techniques such as on board fuel tank inerting or installation of foam in the

tanks would also achieve the goal, but at a cost estimated to be at least

$20 billion over the next 10 years and would be very difficult to retrofit in

current airplanes. Ground inerting, wherein specific tanks are made inert

before flight, at specific airports, is an option that needs future study to

determine; (a) the logistical costs of such a system, and (b) if retrofit

installation of the distribution system internal to the airplane could be achieved

in a cost effective manner.

ï The Working Group considered several concepts that were determined to be

insufficiently advanced technically at this time, for transport airplane fuel tank

use. These included ullage sweeping and explosion suppression systems. 

413 As previously discussed, the FAA subsequently determined that 35 percent was a more accurate

estimate.
Factual Information 217 Aircraft Accident Report

The FTHWG made the following recommendations to the FAA and JAA: 

ï Adopt the proposed new regulatory action on new aircraft designs that would

limit the development of flammable conditions in the fuel tanks, based on the

intended fuel types, to less than 7 percent of the expected fleet operational

time, or provide means to mitigate the effects of an ignition of fuel vapors

within fuel tanks such that any damage caused by an ignition will not prevent

continued safe flight and landing.

ï Continue to investigate means to achieve a cost-effective reduction in

flammability exposure for the in-service fleet and newly manufactured aircraft. 

ï Pursue the studies associated with directed ventilation and ground-based

inerting systems to improve their cost effectiveness.

ï If a practical means of achieving a cost effective reduction in flammability

exposure can be found for the in-service fleet, either at the level specified in

the rule or at some intermediate level (recommendations 2 and 3 above),

consider application of that solution, in combination with other actions (e.g.,

SFAR).

ï If a practical means of achieving a cost effective reduction in flammability

exposure can be found for newly manufactured aircraft, either at the level

specified in the rule or at some intermediate level (recommendations 2 and 3

above), consider application of that solution, in combination with other actions

(e.g., SFAR).[414]

1.18.3.4.3 Notice of Proposed Rulemaking 99-18 (Fuel Tank System Design 

Review, Flammability Reduction, and Maintenance/Inspection 

Requirements)

On October 29, 1999, the FAA published NPRM No. 99-18,415 ìTransport

Airplane Fuel Tank System Design Review, Flammability Reduction, and Maintenance

and Inspection Requirements.î The NPRM indicated that the TWA flight 800 accident and

resulting Safety Board recommendations had prompted the FAA to examine the

414 One FTHWG representative, the representative from the ACAP, submitted a letter, dated July 23,

1998, expressing his dissenting view regarding the final report and recommendations prepared by the

FTHWG. He stated, in part, ìI opposed acceptance of the Report and its recommendations (a) because some

of its key factual findings regarding costs of fixing the problem are unsupported and apparently flawed, (b)

because its recommended rule changes are wholly inadequate to eliminate or significantly reduce the risk of

fuel tank explosions and fires in the near or mid-term future, and (c) because I strongly disagree with the

policy positionÖthat a cost benefit analysis which found that doing nothing for the existing air carrier fleet

would cause the equivalent of three TWA 800 disasters over ten yearsÖî was acceptable. 

In his letter, the ACAP representative further indicated that the ARAC/FTHWG found that a

significant danger exists for fuel tank explosions, especially for CWTs with adjacent heat sources. He stated

that the ARAC/FTHWGís final report indicated that ìthe best available technology to eliminate the risk of

fuel tank explosions is in flight nitrogen inerting which regardless of the ignition source virtually eliminates

the risk and has been successfully used for years on military [transport airplanes]Ö.î 

415 For additional information, see NPRM, ìTransport Airplane Fuel Tank System Design Review,

Flammability Reduction, and Maintenance and Inspection Requirements.î Notice No. 99-18. 64 FR 586644

(October 29, 1999). 
Factual Information 218 Aircraft Accident Report

underlying safety issues surrounding fuel tank explosions, the adequacy of existing

regulations, the service history of airplanes certificated to those regulations, and existing

fuel system maintenance practices. The NPRM proposed to amend the current regulations

to address prevention of ignition sources and minimization of flammable vapors in fuel

tanks in future airplane designs. 

Prevention of Ignition Sources

According to NPRM 99-18, the regulatory authorities and aviation industry have

always presumed that a flammable fuel/air mixture exists in fuel tanks at all times and

have adopted the philosophy that the best way to ensure airplane fuel tank safety is to

preclude ignition sources within fuel tanks. This philosophy has been based on the

application of fail-safe design requirements to the airplane fuel tank system to preclude

ignition sources from being present in fuel tanks when component failures, malfunctions,

or lightning encounters occur. The NPRM noted that airplane fail-safe requirements are

provided in 14 CFR Sections 25.901 (c) and 25.1309416 and that, in general, the FAAís

policy has been to require applicants to assume the presence of foreseeable latent

(undetected) failure conditions when demonstrating that subsequent single failures will

not jeopardize the safe operation of the airplane. However, the NPRM noted that the

requirements of these sections ìhave not been consistently applied and documented when

showing that ignition sources are precluded from transport category airplane fuel tanks.î

The FAA also reviewed SDRs for the transport-category airplane fleet, evaluated

the certification and design practices used on previously certificated airplanes, initiated an

inspection of fuel tanks on existing 747 airplanes, and evaluated the results of an

industry-led inspection of a representative sampling of transport-category airplanes. The

FAA indicated that inspections revealed debris (including lockwire, rivets, and metal

shavings) within the fuel tanks and in scavenge pumps of in-service 747 airplanes.

Further, the FAA observed corrosion and damage to insulation on FQIS probe wiring on

six out of eight probes removed from in-service airplanes. Inspection of fuel tank system

components from out-of-service 747 airplanes revealed damaged (worn, chafed) wiring

and a buildup of conductive sulfide deposits on the FQIS wiring.417 On the basis of its

work, the FAA identified the following conditions as possible contributors to the

development of ignition sources within the fuel tanks: aging of fuel tank system

components, contamination (debris), corrosion of components, and sulfide deposits on

components. 

NPRM 99-18 noted that the intent of existing inspection and maintenance

programs is to ensure that the inherent level of safety that was originally designed into the

system is maintained. The NPRM further noted that historically, FAA-required inspection

programs for fuel tank systems included operational checks, functional checks following

416 These sections are quoted in section 1.16.6.5, in connection with Boeingís CWT ignition fault tree

analysis and report.

417 As previously discussed, tests conducted by the Safety Board during this investigation indicated that

conductive deposits or damaged wiring may result in a location where arcing could occur if high-power

electrical energy was transmitted to the FQIS wiring from another airplane source.
Factual Information 219 Aircraft Accident Report

maintenance actions, overhaul of certain components to prevent dispatch delays, and

general zonal visual inspections conducted concurrently with other maintenance actions.

However, the NPRM noted that specific maintenance instructions to detect and correct

conditions that degrade fail-safe capabilities had not been deemed necessary because it

had been assumed that the original fail-safe capabilities would not be degraded in service.

NPRM 99-18 stated that typical transport-category airplane fuel tank systems are

designed with redundancy and fault indication features such that single component

failures do not result in any significant reduction in safety. Therefore, fuel tank systems

historically have not had any life-limited components or specific detailed inspection

requirements, unless mandated by an AD. Most of the components are inspected ìon

condition,î meaning that some test, check, or other inspection is performed to determine

continued serviceability, and maintenance is performed only if the inspection identifies a

condition requiring correction. Visual inspection of fuel tank system components is by far

the predominant method of inspection for components such as boost pumps, fuel lines,

couplings, wiring. Typically these inspections are conducted concurrently with zonal

inspections or internal or external fuel tank structural inspections. These inspections

normally do not provide information regarding the continued serviceability of components

within the fuel tank system unless the visual inspection indicates a potential problem area.

For example, certain degraded fuel tank system conditions, such as worn wiring routed

through conduit to fuel pumps, debris inside fuel pumps, corrosion to bonding wire

interfaces, would not be detected without dedicated intrusive inspections that are much

more extensive than those normally conducted.418

After its review of the 747 design, the FAA determined that during initial

certification of the fuel tank system, the degree of tank contamination and the significance

of certain failure modes of fuel tank system components had not been considered to the

extent that recent events/service experience indicated was needed. For example, during

initial certification, the manufacturer had demonstrated that the FQIS had been shown to

preclude arcing if FQIS wiring came in contact with the highest electrical voltage

produced on the airplane in the absence of contamination. However, testing conducted

during this investigation has shown that if contamination, such as conductive debris (for

example, lockwire, nuts, bolts, steel wool, corrosion, sulfide deposits, and metal filings),

is located within gaps in the fuel probe, the voltage needed to create an arc is within values

that could result from an electrical short-circuit or current induced in the FQIS probe

wiring by adjacent wiring. According to NPRM 99-18, ìif any of these anomalies were

combined with a subsequent failure within the electrical system that creates an electrical

short, or if [high-intensity radiated fields, or HIRF] or electrical current flow in adjacent

wiring induces EMI voltage in the FQIS wiring, sufficient energy could enter the fuel tank

and cause an ignition source within the tank.î 

On the basis of its review of SDRs and accident and incident reports, the FAA

identified numerous fuel tank system design features, malfunctions, failures, and

418 For more information regarding TWAís visual inspections, see section 1.6.3.1.1.
Factual Information 220 Aircraft Accident Report

maintenance-related actions that would result in a degradation of the systemsí safety

features, including the following: 

Fuel pump-related items

ï ingestion of components (the pump inlet check valve or the pump inducer) in

the pump impeller, resulting in debris in the fuel tank;

ï stator winding failures during operation of the fuel pump, resulting in arcing

through the fuel pump housing;

ï inappropriate wrapping of the fuel pump windings, resulting in deactivation of

the pumpís thermal protective features; 

ï absence of cooling port tubes between the pump assembly and the pump motor

(has occurred as a result of improper fuel pump overhaul);

ï extended fuel pump operation in empty fuel tanks (dry-running fuel pumps); 

ï use of steel impellers that could produce sparks if debris enters the fuel pump; 

ï exposure of electrical connections within the fuel pump housing with

inadequate clearance to the fuel pump cover, resulting in arcing; 

ï internal wires contacting the pump rotating group, energizing the rotor and

causing arcing at the impeller/adapter interfaces;

ï poor bonding of components to structure and/or across component interfaces; 

ï insufficient ground fault current protection capability; 

Wiring inside fuel tanks

ï chafing/wearing of Teflon sleeve and wire insulation, allowing arcing through

metallic conduits into fuel tanks;

Fuel pump connectors

ï electrical arcing at connections within electrical connectors because of bent

pins or corrosion;

ï corrosion of electrical connectors in the fuel pump motor, resulting in fuel

leakage and subsequent fuel-fed fire outside the fuel tank and electrical arcing

through the connector housing (outside the fuel tank); 

ï improper materials in connector design;

FQIS wiring

ï degradation of wire insulation (cracking), corrosion, and sulfide deposits at

electrical connectors; 

ï unshielded FQIS wires routed in wire bundles with high-voltage wires. 
Factual Information 221 Aircraft Accident Report

FQIS probes

ï corrosion and sulfide deposits, resulting in reduced breakdown voltage in FQIS

wiring;

ï terminal block wiring clamp features at electrical connections on fuel probes,

damaging wire insulation;

ï contamination in fuel tanks, causing reduced arc path between FQIS probe

walls; 

Bonding straps

ï corrosion to bonding straps;

ï loose or improperly grounded attachment points;

ï static bonds on fuel tank system plumbing connections inside the fuel tank

worn by wing movement and/or corrosion;

Electrostatic charge

ï use of nonconductive reticulated polyurethane foam that holds electrostatic

charge buildup; and

ï spraying of fuel into fuel tanks through inappropriately designed refueling

nozzles or pump cooling flow return methods. 

NPRM 99-18 stated that the FAAís ìreview of service history, design features, and

maintenance instructions of the transport airplane fleet indicated that aging of fuel tank

system components and unforeseen fuel tank system failures and malfunctions have

become a safety issue for the fleet of turbine-powered transport category airplanes.î

NPRM 99-18 proposed an SFAR that would require transport-category airplane

type certificate (TC) holders for transport-category airplanes and holders of supplemental

type certificates (STC) that affect the transport-category airplaneís fuel system419 to

conduct a safety review of the fuel tank system designed to show that fuel tank fires or

explosions will not occur. Specifically, the TC or STC holder would be required to

determine whether the fuel system design meets the existing requirements of 14 CFR

Section 25.901420 and the requirements of Section 25.981(a) and (b), which would include

several provisions added by the NPRM. If the design does not meet these requirements,

419 The proposed rule would apply to airplanes that had TCs issued after January 1, 1958, and that have a

certificated passenger capacity of at least 30 people or payloads of 7,500 pounds or more and proposed a

12-month compliance period for the review and an 18-month compliance period for the maintenance and

inspection instructions development. The FAA has determined that 35 TCs and 168 STCs would need a fuel

tank system design assessment.

420 Section 25.901 requires, in part, that major components of airplane powerplant installations

(including fuel tanks) be electrically bonded to the other parts of the airplane and that for each powerplant

installation, no single failure or malfunction or probable combination of failures will jeopardize the safe

operation of the airplane. 
Factual Information 222 Aircraft Accident Report

the SFAR would require the TC or STC holder to develop the necessary corrective design

changes. 

As previously discussed, Section 25.981 currently requires that ìa safe marginî

exist between the temperature at any place inside a fuel tank where fuel ignition is

possible and the lowest expected auto ignition temperature of the fuel in the fuel tank.

NPRM 99-18 proposed to revise Section 25.981 so that subsection (a) would prohibit an

ignition source from being present at any point in the fuel tank system ìwhere catastrophic

failure could occur due to ignition of fuel or vapors.î The new rule would require that this

be shown by demonstrating (1) compliance with the existing requirement regarding auto

ignition temperature and (2) that an ignition source in the fuel tank system could not result

from any single failure, from any single failure in combination with any latent failure

condition not shown to be extremely remote, or from any combination of failures not

shown to be extremely improbable.421 The revised Section 25.981 would also require, in

subsection (b), that critical design configuration control limitations, inspections, or other

procedures be established as necessary to prevent development of ignition sources.

According to NPRM 99-18, the TC and STC design approval holder would be

expected to do the following to comply with the SFAR:

develop a failure modes and effects analysis (FMEA) for all components in the

fuel tank system. Analysis of the FMEA would then be used to determine whether

single failures, alone or in combination with foreseeable latent failures, could

cause an ignition source to exist in a fuel tank. A subsequent quantitative fault tree

analysis should then be developed to determine whether combinations of failures

expected to occur in the life of the affected fleet could cause an ignition source to

exist in a fuel system.[422]

The SFAR would also require TC and STC holders to develop all maintenance and

inspection instructions necessary to maintain the design features required to preclude the

existence or development of an ignition source within the fuel tank system.423 According

to NPRM 99-18, the FAA anticipates that the safety review would identify critical areas of

the fuel tank and other related systems that would require maintenance actions to account

for the effects of aging, wear, corrosion, and possible contamination of the fuel tank

system. The NPRM noted, for example, that service history indicates that copper-sulfide

421 According to the NPRM, for the purposes of the new rule, extremely remote failure conditions are

those not anticipated to occur to each airplane during its total life but that may occur a few times when

considering the total operational life of all airplanes of the type. Extremely improbable failure conditions are

those not anticipated to occur during the entire operational life of all airplanes of one type. 

422 The Safety Board identified numerous potential ignition hazards (including, but not limited to, aging

components, contamination and corrosion of components, and sulfide deposits on components) during the

TWA flight 800 investigation. The NPRM stated that these failure conditions and any other foreseeable

failures should be assumed when performing the FMEA analysis. However, the NPRM indicated that

uncontained engine failures; terrorist actions, such as bombs; and the effects of postcrash or other fires from

outside the tank would be outside the purview of the proposed FMEAs. 

423 The SFAR also states that ìno person may operate a turbine-powered transport category

aircraftÖunless instructions for maintenance and inspection of the fuel tank system are incorporated into its

inspection program.î 
Factual Information 223 Aircraft Accident Report

deposits may form on fuel tank components and that this could degrade the intended

design capabilities by providing a possible arcing mechanism. Therefore, the NPRM

noted that it might be necessary to provide maintenance instructions to identify and

eliminate such deposits.

Minimizing Development of Flammable Vapors in Fuel Tanks 

NPRM 99-18 noted that, in addition to its review of potential fuel tank ignition

sources, the FAA had ìundertaken a parallel effort to address the threat of fuel tank

explosions by eliminating or significantly reducing the presence of explosive fuel air

mixtures within fuel tanks of new type designs, in-production [airplanes], and the existing

fleet of transport airplanes.î According to the NPRM, the FAA agreed with the intent of

the regulatory text recommended by the FTHWG. However, because no economic

evaluation had been completed to determine if practical means were available to reduce

the exposure below the 7 percent of fleetwide operational time proposed by the FTHWG,

the NPRM stated that the FAA was proposing a more objective regulation intended to

minimize exposure to operation with flammable conditions in fuel tanks.

NPRM 99-18 proposed to add a new subsection (c) to 14 CFR Section 25.981,

which would require that fuel tank installations in newly designed airplanes include a

means to minimize the development of flammable vapors in fuel tanks, or to mitigate the

effects of an ignition of fuel vapors within fuel tanks, such that no damage caused by an

ignition will prevent continued safe flight and landing. According to the NPRM, the

proposal was not intended to prevent the development of flammable vapors because total

prevention has not been found to be feasible; instead, the proposal is intended as an

interim measure to preclude, in new designs, the use of design methods that result in a

relatively high likelihood that flammable vapors will develop in fuel tanks. (The NPRM

noted that certain design methods, such as the location of air conditioning packs directly

below a fuel tank without provisions to reduce transfer of heat from the packs to the tank,

increase the likelihood that flammable vapors will develop in fuel tanks.) The NPRM

further stated that the intent of the proposal is to require that fuel tanks are not heated and

that they cool at a rate equivalent to that of a wing tank in the transport-category airplane

being evaluated. 

Draft Advisory Circulars

In connection with regulatory changes proposed in NPRM 99-18, on January 12,

2000, the FAA issued two draft ACs. FAA Draft AC 25.981-1X, ìFuel Tank Ignition

Source Prevention Guidelines,î ìprovides guidance for demonstrating compliance with

the [proposed] certification requirements for prevention of ignition sources within the fuel

tanks of transport category airplanes.î The draft AC states that ì[p]rior to conducting a

fuel system safety assessment, each applicant should assemble and review relevant lessons

learned from their previous products, their suppliers, and any other available sources, to

assist in identifying any unforeseen failures, wear, or other conditions that could result in

an ignition source.î424 The draft AC provides detailed guidance about potential fuel tank

ignition sources, accepted design practices, ignition source failure analysis, electrical

sources, friction sparks, surface ignition, and instructions for continued airworthiness.
Factual Information 224 Aircraft Accident Report

Further, the draft AC states that the ignition source failure analyses required by the

proposed SFAR

are intended to address system failures that may result in the presence of an

ignition source in the fuel tanks. These regulations are not intended to address

failures or conditions that could lead to ignition of fuel vapors from such sources

as uncontained engine debris, external fires following engine separation or failure,

damage resulting from explosive materials, such as bombs, or post crash fire

heating of tank surfaces or propagation of fire through the airplane vent system

into the fuel tanks. These hazards are addressed by other regulations. 

FAA Draft AC 25.981-2X, ìFuel Tank Flammability Minimization,î ìprovides

information and guidance concerning compliance with the [proposed] airworthiness

standards for transport category airplanes pertaining to minimizing the formation or

mitigation of hazards from flammable fuel air mixtures within fuel tanks.î The draft AC

reiterates, as noted in NPRM 99-18, that the intent of the proposed regulation is to require

that the exposure to formation or presence of flammable vapors is equivalent to that of an

unheated wing tank in the transport-category airplane being evaluated. The draft AC

contains detailed guidance about methods for minimizing fuel tank flammability and

acceptable means to mitigate the effects of an explosion. 

Safety Board Comments on NPRM 99-18

In a March 23, 2000, letter, the Safety Board submitted its comments on

NPRM 99-18. The Board stated that it generally supported the proposed regulatory

changes but urged the FAA to give further consideration to some issues. Specifically, with

regard to the proposed requirement for FMEAs, the letter stated that the ìSafety Board

generally supports the intent of the SFAR in requiring an FMEA analysis, but is concerned

about the construction of the FMEAs, as well as the thoroughness and integrity of the data

that will be used.î The Board cited the concerns expressed by Safety Board investigators

and NASA specialists about the Boeing fault tree analysis report submitted to the Board in

connection with the TWA flight 800 investigation.425 The Board further stated the

following:

the Safety Boardís concerns about the FMEAs are amplified by the fact that no

single source exists for reliable and comprehensive data on component failures or

malfunctions. Because the calculations in an FMEA are based on failure rates,

incomplete or inappropriate failure data can skew the results of an examination.

The Board is aware that service history data maintained by manufacturers do not

capture data from all operators. Further, the Board has found that the amount of

data provided by the manufacturers of replacement component parts sometimes

greatly exceeds the data provided by the aircraft manufacturers (possibly because

replacement parts suppliers can sell parts directly to operators and repair

facilities). Although the FAA collects a significant amount of data about

mechanical failures through its Service Difficulty Report (SDR) program, even

424 The draft AC also states that ìservice history has shown that ignition sources have developed in

airplane fuel tanks due to unforeseen failure modes or factors that were not considered at the time of original

certification of the airplane, including sparks, or hot surfaces within the fuel tanks.î

425 For more information about this fault tree analysis report, see section 1.16.6.5.
Factual Information 225 Aircraft Accident Report

these data are incomplete.[426] Other sources of potentially relevant data are the

service histories maintained by the military of its variants of commercial airliners

and the Boardís accident and incident investigation database; however, neither of

these sources provides complete data either. 

Further, the many affected TC and STC holders (some of which are not the

original designers or manufacturers) may have varying levels of experience with

developing FMEAs. In addition, the Safety Board is concerned that engineers

working for TC and STC holders may not recognize the existence or significance

of certain hazards and that potentially competing interests may affect the quality

and thoroughness of some FMEAs. In cases in which the TC or STC holder no

longer exists, FAA personnel with varying levels of skill and experience may have

to conduct the FMEAs. Finally, the Board is concerned that the FAA may have an

insufficient number of staff who are trained to properly evaluate an FMEA. 

Therefore, to ensure the integrity and effectiveness of the fuel tank system safety

review, the Safety Board urges the FAA to develop and provide adequate

standards and criteria to guide the development of the FMEAs and fault tree

analyses. In particular, because there is no single comprehensive collection system

that contains data on the failure of airplane components and because of the

inadequacies that exist in each source of data, those guidelines should specify that

the data used for the FMEAs must be collected from all available sources,

including operators, manufacturers, and appropriate government agencies.

The FAA needs to provide adequate oversight and auditing of the FMEA analysis

results to ensure their technical accuracy and integrity. In particular, such

oversight and auditing should include a review of the data sources used to ensure

that all available and appropriate sources of failure data have been taken into

account. FAA oversight and auditing can also provide a method to identify

potential deficiencies in the FMEAs that might be overlooked or not recognized

by engineers and designated engineering representatives who work with these

systems on a daily basis. 

Regarding the proposed requirement for maintenance and inspection instructions,

the Safety Board indicated that it

strongly endorses continuing airworthiness through improvements to

maintenance, inspection, and minimum equipment lists. However, given the very

general nature of some current inspection criteria pertaining to fuel tank safety (as

documented by the FAA in its Transport Non-Structural Systems Plan), the Board

is concerned that the instructions resulting from the SFAR requirement may be

similarly broad and, therefore, potentially ineffective. Many potential ignition

sources (such as hidden cracks in wiring, sulfide deposits, and use of

inappropriate materials) may not be apparent during a general visual inspection.

Therefore, the Board urges the FAA to ensure that the maintenance and inspection

instructions developed as a result of this SFAR are detailed and specific enough to

provide mechanics with useful inspection criteria and to ensure that they are

properly trained about how to effectively carry out those instructions, including a

requirement for a detailed inspection of each component in any area that may be

exposed to fuel or fuel vapors. 

426 For additional information regarding the Safety Boardís concerns about and recommendations made

pertaining to the SDR program, see section 1.18.2.1. 
Factual Information 226 Aircraft Accident Report

Regarding the regulatory proposal to minimize development of flammable fuel

vapors in new designs, the Safety Board indicated that it was ìconcerned that the NPRM

does not propose any regulatory changes that address fuel tank flammability in current

designs and in the existing fleetî and that ì[t]his is especially disturbing because some

operational measures (such as limiting the on-ground operating time of air conditioning

packs and substituting a ground-based cool air supply and cooling or ventilating the pack

bay) that can reduce current levels of flammable vapors could be accomplished

immediately.î The Board stated that it was ìaware that the FAA is conducting research

into on-ground fuel tank inerting systems for the existing fleet,[427 and that because] the

Board believes that fuel tank inerting is a promising, near-term method that could

dramatically reduce fuel tank flammability in the existing fleet, it strongly supports the

FAAís continued work in this regard and looks forward to regulatory implementation.î

Finally, the Safety Board stated the following:

[a]lthough FAA staff have indicated to Safety Board staff that the FAA does not

intend to endorse the FTHWGís proposed exposure criteria [of 7 percent of the

fleet operational time], the Board nonetheless wishes to register its concerns about

those criteria. Because it is a fleetwide average, it does not account for increased

risks that may exist at specific locations, during certain time periods, or for certain

flights.[428]

 In addition, the premise that transport airplane fuel tanks located in the

wings have an acceptable safety record is unacceptable because wing fuel tanks

have exploded.[429]

 The Safety Board believes that the goal should be to

completely eliminate the development of flammable vapors in fuel tanks to the

greatest extent technically feasible (such as would result from the use of onground inerting systems). 

1.18.3.4.4 FAAís Aging Transport Non-Structural Systems Plan

On July 25, 1996, White House Commission on Aviation Safety and Security

(WHCASS) was chartered. The WHCASS subsequently recommended that the FAA

expand its aging aircraft program to cover nonstructural systems by (1) conducting full

and complete teardowns of selected aircraft scheduled to go out of service, (2) establishing

a ìlead-the-fleetî research program, (3) expanding the FAA-DoD-NASA cooperative

aging aircraft program, (4) expanding programs of the Airworthiness Assurance Working

Group to include nonstructural components, and (5) encouraging the development of

modern technical means to ensure and predict the continued airworthiness of aging

nonstructual systems. 

427 The Safety Board noted that the FAA is also engaged in research regarding on-board inerting systems

for future designs. 

428 The Safety Board noted that of the 16 fuel tank explosions documented by the ARAC, 11 occurred in

the months of May, June, and July, and only 2 occurred outside of warm climates or summer months.

Further, short flights in warm climates can have a dramatically higher exposure time than the fleetwide

average. For these flights, attempting to reduce the flammability of CWTs to the level of wing tanks would

be of little benefit.

429 For example, as previously indicated, the Board noted that wing tank explosions caused the May 9,

1976, accident involving a 747 operated by the Iranian Air Force near Madrid, Spain, and the January 15,

1972, accident involving a U.S. Navy C-130 near Cambridge, Maryland.
Factual Information 227 Aircraft Accident Report

The FAA participated in the Safety Boardís inspections 26 airplanes in

January 1998, then inspected additional airplanes by the end of March 1998. The FAAís

examinations revealed deterioration of wiring components (wires, bundles, connectors,

grounds, clamps, shielding), contamination of wire bundles with excessive dust and fluids,

cracking of outer layers of multilayer insulation, ground terminals with resistance

measurements outside specifications, corrosion on connector pins, and improper wire

installations/repairs. The FAA also found that aging wire becomes stiff and easily cracked;

wire bundles are often difficult to inspect; and contamination of wire bundles with metal

shavings during maintenance is a common occurrence. In response to these findings and

the WHCASSís recommendation, in July 1998, the FAA issued the Aging Transport

Non-Structural Systems Plan.

The FAAís plan concluded that current maintenance practices do not adequately

address wiring components; inspection criteria are too general (for example, ìperform a

general visual inspectionî); unacceptable conditions are not described in sufficient detail;

repair instructions/data are hard to locate and extract from Boeingís SWPM, as currently

presented; wire replacement criteria may not be adequate; and current

incident/maintenance reporting procedures do not allow for easy identification of failures. 

The FAA then identified the following tasks:

ï establish an Aging Transport Systems Oversight Committee to coordinate the

various aging systems initiatives within the FAA;

ï conduct an in-depth review of the aging transport fleet (including service

history) and make model-specific safety recommendations related to airplane

systems;

ï enhance airplane maintenance to better address aging airplane systems;

ï improve inspection criteria for wiring (need more complete descriptions of

undesirable conditions);

ï define acceptance criteria for corrosion on flight control actuators, associated

linkages, and hydraulic fittings;

ï revise [Maintenance Symposium Group] MSG-3 process to address

catastrophic events associated with wire failures as MSG-3 review items and

identify maintenance tasks/intervals to address;

ï establish a means to minimize contamination of wiring from metal shavings

(vacuuming is not sufficient, can cause shorting of wires and failure of

electrical systems, may become a serious problem for new airplanes where

0.008-inch insulation thickness wire is used versus 0.020-inch insulation in

current aging airplanes);

ï review inspection/repair training to ensure it adequately addresses recognition

and repair of aging wiring components;

ï develop guidance material to address electromagnetic compatibility, lightning,

and HIRF protection in maintenance programs;

ï review and update Chapter 20 of manufacturerís standard practices of wiring;
Factual Information 228 Aircraft Accident Report

ï add aging systems tasks to the aging airplane research program, determine if a

service life for airplane wire is appropriate, establish condition of aging wiring

components and validate the adequacy of visual inspections, develop

nondestructive testing tools for inspection of wiring systems,430 establish aging

effects on lightning and HIRF protection systems, develop arc fault circuit

breaker, and perform destructive testing of flight control linkages;

ï improve reporting of accident/incident and maintenance actions involving

wiring system components (wires, connectors, wire shields, grounds, circuit

breakers), urge ATA to establish appropriate codes, improve formats, add data

to National Aviation Safety Data Analysis Center databases, and recommend

that aging components removed during maintenance be examined for safety

implication of failure; 

ï evaluate the need for additional maintenance of transport-category airplane

fuel system wiring and address any unsafe conditions, review fleetwide service

problems to identify any unsafe fuel system wiring conditions, and develop an

SFAR to address fuel system wiring design and maintenance practices; and

ï improve wiring installation drawings and instructions for continuing

airworthiness, task SAE to write an aviation recommended practice to define

best practices for modification of existing wiring on transport-category

airplanes, develop training aids in wiring installation practices for certification

engineers and designees, develop job aid for certification engineers and

designees to use in evaluation of adequacy of wire systems installation

drawings and instructions for continuing airworthiness. 

In November 1998, the FAA established the ATSRAC, and on January 19, 1999,

defined its charter in FAA Order 1110.127. According to the charter, the ATSRACís task

is to propose FARs revisions and associated guidance material as appropriate to ensure

that nonstructural systems in transport-category airplanes are designed, maintained, and

modified in a way that ensures their continued operational safety throughout the service

life of the airplanes. Its duties are to recommend standardization clarification and to

upgrade regulations, standards, and procedures related to aging systems as necessary to

institutionalize lessons learned from aging systems review and recommend mandatory

retroactive rules (ADs) to ensure safety of the transport-category airplane fleet. 

On January 11, 1999, the FAA issued ìTerms of Reference,î which identified the

following five major tasks for the ATSRAC: (1) conduct an in-depth survey of the

condition of aging transport-category airplane fleetwide systems and propose

model-specific safety recommendations related to airplane systems that will eliminate or

significantly reduce the hazards associated with the types of age-related degradation

430 In this regard, the FAA plan noted that ì[w]ire is difficult to inspect in several areas of the airplane

due to inaccessibility. Wire inside conduits is not inspectable by visual means. These difficulties, associated

with the inspection of aging wiring bundles and connectors, can be overcome by the development and use of

non-destructive testing tools to assess the state of the wiring in an airplane. The FAA will explore what tools

already exist for these purposes. Where tools do not exist, promising technologies will be explored for

adoption to the problems associated with inspection of aging wiring systems.î 
Factual Information 229 Aircraft Accident Report

displayed by the fleet; (2) review the service history, manufacturerís SBs and SLs, and

applicable ADs for the aging transport-category airplane fleet and identify information

that pertains to aging systems for possible mandatory action; (3) improve general

maintenance criteria for airplane systems to ensure that aging systems-related problems

are identified and corrected; (4) review, update, and simplify manufacturerís standard

practices for wiring; and (5) review air carrier and repair station training programs for

nonstructural systems inspection and repair to ensure that they adequately address aging

wiring system components and other nonstructural systems.

In connection with improvement of maintenance criteria, the FAA noted that

ìmaintenance procedures currently in use in the air transport industry may not adequately

or proactively address aging non-structural systems.î The FAA specified several subtasks

associated with this task, including defining improved inspection criteria for wiring,

connectors, and associated components; defining practices to eliminate wire bundle

contamination (with metal shavings, fluids, or other damaging materials) during

maintenance; defining acceptance criteria for corrosion of systems components; and

proposing a process to ensure that components removed during maintenance are examined

for safety implications of the observed failures and the results are tracked to the original

service problem. 

1.18.4 Witness Information

1.18.4.1 General Witness Information

In the days immediately following the accident, the Safety Board announced its

intention to form a Witness Group to interview witnesses to the accident. However, the

Board deferred to the concerns expressed by Federal criminal investigators regarding the

need for confidential treatment of eyewitness information while the criminal investigation

was ongoing. Hence, the Board did not initially share witness information with private

sector parties and did not use its party system to conduct witness interviews in its

traditional fashion. Instead, a Safety Board investigator reviewed FBI interview notes and

briefed other Board investigators on their contents.

In November 1996, the FBI agreed to provide the Safety Board with access to

witness documents in which personally identifying information had been blacked-out

(redacted). A Witness Group was formed to review these documents and conduct a limited

number of witness interviews in cooperation with the FBI.431 Upon completion of these

interviews and the groupís review of the redacted witness documents, this Witness Group

was disbanded. 

The FBI concluded its active criminal investigation of the crash in

November 1997,432 and in April 1998, it provided redacted witness documents433 to Safety

Board investigators. The Board formed a second Witness Group to study the documents

431 The Safety Board intended to present information about the witnesses at its December 1997 public

hearing; however, the FBI objected, and the Board did not address this topic at its hearing.
Factual Information 230 Aircraft Accident Report

and prepare them for inclusion in the public docket. (Unless otherwise specified, all

subsequent references to the ìWitness Groupî are to the second Witness Group.) 

The witness documents were summaries of information434 collected by FBI agents

during interviews with potential witnesses. The Witness Group determined that the

documents contained 736 witness accounts.435 Because group members believed that it

was not necessary and because of the amount of time that had elapsed (about 21 months)

when the Safety Board received information about the identity of the witnesses,436 the

Witness Group chose not to reinterview most of the witnesses;437 instead, the group chose

to rely on the original FBI documents as the best available evidence of the observations

initially reported by the witnesses.

According to these documents, 670 witnesses reported seeing something that the

Witness Group determined was probably related to the accident, and nearly 250 reported

hearing something that was determined to be likely related to the accident.438 Figure 40 is

a map that shows the witness locations and the location of the accident site.

The Witness Group determined that of the 736 witnesses, 239 were sound

witnesses,439 179 were sight and sound witnesses, 258 were streak-of-light440 witnesses,

432 Although the FBI ceased its active investigative efforts in November 1997, it indicated that it

remained prepared to resume the investigation at any time, should information indicate that such resumption

was warranted. 

433 The FBI provided the Safety Board with witness documents that the Board could release in the public

docket; to this end, FBI personnel removed personal information that would have identified the witnesses

from the documents. For additional information, see the Witness Group Chairmanís Factual Report, dated

February 9, 2000. 

434 The FBI stated that no verbatim records of the witness interviews were produced. The summaries

provided to the Safety Board were generally written in the words of the agents who conducted the interviews

and not necessarily in the words of the witnesses themselves. Witnesses were not asked to review or correct

the documents. 

435 A witness was defined by the Witness Group as ìanyone who reported hearing a sound and/or seeing

an event or object or objects (including smoke or fire) in the sky in the general vicinity of the accident site,

on July 17, 1996, at the approximate time of the accident.î

436 Research indicates that the most accurate witness information is obtained immediately after an event

is observed. See section 1.18.4.3 for additional information regarding the effects of time on witness

recollections. 

437 One exception was the captain of Eastwind flight 507, who, because of his vantage point relative to

TWA flight 800 and his experience as an airline pilot, was reinterviewed by the Witness Group on March 25,

1999. 

438 For additional information, see the Witness Group study report, dated February 9, 2000. 

439 The documents for these witnesses reported that they heard sounds such as a boom, a roar, or a

thunderlike rumble. Two hundred seventy-five of the witness documents pertaining to the 670 sight

witnesses did not mention hearing any sounds, and 216 of the remaining sight witnesses did not hear

anything that the Witness Group judged to be related to the accident. 

440 The Witness Group defined a streak of light as ìan object moving in the sky that could be variously

described in witness documents as a point of light, fireworks, a flare, a shooting star, or something similar.î

The definition noted that a streak of light was usually described as ascending, but could also be described as

arcing over and/or descending. The Witness Group agreed that the streak of light definition must not meet

the definition of a fireball and that it must precede any report of a fireball by that witness.
Figure 40. Map showing the witness locations and the location of the accident site.
Factual Information 232 Aircraft Accident Report

and 599 were fireball witnesses. Of the 258 streak-of-light witnesses, 38 reported that the

streak was ascending vertically or nearly so,441 18 indicated that it originated from the

surface of the earth, and 7 reported that it originated at the horizon.442 Of the 599 fireball

witnesses, 264 reported seeing the fireball originate, 200 reported seeing the fireball split

into two fireballs, and 217 reported observing the fireball hit the surface of the water (or

disappear below the horizon). In addition, 210 witnesses reported seeing both a fireball

and a streak of light.

Sufficient information was present in the materials provided by the FBI for the

Witness Group to plot the reported position (approximate latitude and longitude) of

591 witnesses and the situation (land, sea, or air) of 699 witnesses.443 Of the 591 witnesses

whose reported positions were plotted, seven were positioned within 5-nm ground

distance of the last secondary radar return from TWA flight 800. About 200 witnesses

were positioned between 5- and 10-nm ground distance of the last secondary radar return.

More than 200 witnesses were positioned between 10- and 15-nm ground distance of the

last secondary radar return. About 80 witnesses were positioned between 15- and 20-nm

ground distance of the last secondary radar return. Nearly 100 witnesses were positioned

more than 20-nm ground distance from the last secondary radar return. 

Of the 699 witnesses whose witness documents contained sufficient information to

determine their reported situation at the time of the accident,444 460 were on land,

202 were on or in the water, and 37 were in aircraft. Most of the witnesses on land were

outside, and most of the witnesses on the water were in boats. Of the 460 witnesses on

land, 172 reported seeing a streak of light,445 and 365 reported seeing a fireball (122 of

these witnesses reported seeing the fireball split into two fireballs during its descent).

441 The Witness Group determined that of the 220 streak-of-light witnesses who did not characterize the

motion of the streak as vertical or nearly so, 126 provided some information about the lateral motion of the

streak. Of these, 45 witnesses reported that the streak moved to the east, 23 reported that it moved to the

west, 18 reported that it moved to the south, and 4 reported that it moved to the north. Thirty-six of the

streak-of-light accounts contained no information about the direction of travel. 

442 Some of the 38 witnesses who reported the streak ascending vertically or nearly so may also have

reported that it originated from the surface of the earth or from the horizon. Thirteen of these 38 witnesses

were in vehicles (including boats), and one was paddling a surfboard at the time of the reported observation.

A total of 56 witnesses reported seeing a streak of light ascending vertically or nearly so and/or a streak of

light that originated from the surface of the earth or the horizon.

443 Six calculations were made for each of the witnesses for whom a position could be determined:

ground distance, slant range, elevation angle, initial azimuth, final azimuth, and sound time. Ground

distance was defined as ìthe linear distance in feet along the surface from the witnessís location to the

surface of the water directly under the last secondary radar return for flight 800.î Slant range was defined as

ìthe elevated distance in feet from the witnessís position and the position of flight 800 at altitude at the time

of the last secondary radar return, calculated using ground distance and the airplaneís altitude.î Elevation

angle was defined as ìthe angle in degrees formed from the witnessís line of sight to flight 800.î The initial

azimuth was defined as ìthe bearing (in degrees true) from the witness to the accident airplane at the time of

the last secondary radar return.î The final azimuth was defined as ìthe bearing in degrees true from the

witness to a point in the northeast corner of the debris field.î The sound time was defined as ìthe

approximate time required for sound to travel to the witness from flight 800ís position at the time of the last

secondary radar return.î

444 The witness document(s) for 2 of the 258 streak-of-light witnesses, 15 of the 599 fireball witnesses,

and 1 of the 200 witnesses that reported seeing the fireball split in two did not contain sufficient information

to determine that witnessís situation at the time of the accident.
Factual Information 233 Aircraft Accident Report

About one-third of the witnesses on land reported hearing something related to the

accident. Of the 202 witnesses who were on or in the water, 78 reported seeing a streak of

light, and 182 reported seeing a fireball (61 of whom reported seeing the fireball split).

About 25 percent of those witnesses on or near the water reported hearing something

related to the accident. All of the 37 airborne witnesses reported seeing a fireball (16 of

whom reported seeing the fireball split in two), and 6 of the 37 reported seeing a streak of

light. 

A witness document pertaining to a professional fisherman who, according to the

document, had been fishing in the area of the accident for about 5 or 6 months preceding

the accident, indicated that there had been ìa lot of flare activityî in the area during that

time. A document pertaining to another witness, who was on board a USCG cutter that

responded to the accident, indicated that by the time the cutter arrived at the accident site

about 30 minutes after the accident, ìfishing boats in the area had already begun to set off

flares.î 446

The Witness Group reached 17 conclusions, including the following:

ï The streak of light reported by many (of the 258 streak-of-light witnesses) may

have been the accident airplane during some stage of its flight447 before the

fireball developed. Most of the streak-of-light accounts are generally

consistent with the calculated flightpath of the accident airplane; however,

38 witnesses described a streak of light rising straight up or nearly so. These

38 witness accounts seem to be inconsistent with the calculated flightpath of

the accident airplane.

ï The descending fireball (observed by 599 witnesses) was the accident airplane.

In all likelihood, the witnesses who watched this fireball split were observing

the final stages of the structural breakup of the aircraft.

1.18.4.2 Witness Document Issues 

According to the Witness Group study report, the Witness Group identified several

aspects of the witness documents that made it difficult to extract accurate and reliable

information from them, including possible interviewer and/or interviewee bias (for

example, the assumption that a missile was used against the airplane), potentially flawed

445 The 38 witnesses who characterized the motion of the streak of light as vertical or nearly so were

found to be widely dispersed throughout Long Island. The closest of these witnesses was situated about

7.8-mn slant range from TWA flight 800, and the farthest was about 47-nm slant range from TWA flight

800.

446 For more information about this witness document, see section 1.18.4.4.

447 Central Intelligence Agency (CIA) analysts, who initially evaluated certain witness information

provided by the FBI, also concluded that the majority of the witness observations must have occurred well

after the initial explosion and, consequently, none of those observations could have been of a missile that

caused the explosion. The CIA analysts concluded that the remaining eyewitness accounts described events

fully consistent with observations expected if only the aircraft in various stages of crippled flight were being

observed. The CIA analysts concluded that there was absolutely no evidence that a missile was employed to

shoot down TWA flight 800.
Factual Information 234 Aircraft Accident Report

distance and direction information, multiple witness reports summarized in a single

account, recording of witness speculations and conclusions, imprecise or vague language,

conflicting information, and errors concerning the origin of the streak of light.

According to an FBI Special Agent involved with the FBIís earlier criminal

investigation who met with the Witness Group on September 30, 1998,448 the FBI initially

suspected that a missile might have been used against TWA flight 800; therefore, FBI

agents tended to use the witness documents to capture information that appeared relevant

to its criminal investigation involving a missile rather than to record a complete

accounting of the events described by the witness. In addition, missile experts were

present at some interviews. According to the Special Agent, FBI agents were provided

with suggested interview questions pertaining to missiles. These questions, which were

included in the documents provided by the FBI pertaining to Witness 32 (and which the

Special Agent indicated were asked in several cases), included the following:

ï ìHow long did the missile fly?î 

ï ìWhat does the terrain around the launch sight look like? Were scorch marks

visible?î

ï ìWhere was the sun in relation to the aircraft and the missile launch point?î

The Witness Group study report cited the presence of missile experts and questions

such as these as potential sources of interviewer or interviewee bias that could influence

the outcome of the interview. Other examples of possible interviewer/interviewee bias

cited in the Witness Group study report included the following:

ï According to a witness document, a witness located about 10.4-nm slant range

from TWA flight 800449 described an ascending red ball. The document further

states, ìUpon impact, [witness' name redacted] observed a large fireball.î

Neither the witness document nor the attached interview notes450 stated that the

witness saw anything other than the red ball in the sky; thus, it is unclear what

the ìimpactî mentioned in the document is. Although this incongruity could be

because of other reasons, the witness appears to have described an ascending

red ball and a large fireball, which the interviewer related using the word

ìimpact.î It is also possible that the witness used the word ìimpact,î but the

original interview notes used to generate the witness document do not contain

the word ìimpact.î

448 On September 30, 1998, an FBI agent who had been substantially involved in the FBIís missile

investigation met with the Witness Group to discuss the witness-related processes used by the FBI during

this investigation. 

449 For more information about individual witnesses, see the Witness Group study report, which contains

witness numbers that can be correlated to witness documents.

450 Interview notes were not available for most witness documents. 
Factual Information 235 Aircraft Accident Report

ï Two other witnesses (both of whom were 8.2-nm slant range from TWA flight

800), described seeing a flarelike object but, according to the document, ìthey

did not see what [the] flare struck, but it exploded in air into a large orange

fireball.î This characterization may suggest that the interviewer and/or the

interviewees believed that the flarelike object was a missile, which must have

struck something.

According to the Witness Group study report, some witness documents noted that

witnesses acknowledged that they did not realize that they had observed an airplane

accident until they saw media accounts of the accident, and some documents specifically

stated that witnesses made conclusions about what they observed after learning about the

accident in the media. For example, one document stated that a witness (located about

13.2-nm slant range from TWA flight 800) did not think anything about his observation

until he learned about the TWA flight 800 accident while watching the evening news.

Another document indicated that a witness located 31-nm slant range from TWA

flight 800 concluded that he had observed a missile after hearing news accounts about the

crash. The Witness Group study report noted that media coverage about the crash and the

associated criminal investigation might have led to bias on the part of some

interviewees.451

Examination of the witness documents also indicated that distance estimates and

direction of travel information varied considerably. For example, two witnesses who were

in the same location observing the accident described seeing an ascending flarelike object.

One witness estimated that the flarelike object was about 30 miles away, whereas the other

witness estimated that it was 5 or 6 miles away. (Subsequent calculations indicated that the

accident site was located about 26 statute miles ground distance away from their location.)

Additionally, in general, there were notable variations in the witness accounts regarding

the direction of motion of the streak of lightóas mentioned in section 1.18.4.1,

38 documents reported that the streak of light moved vertically, 45 reported that it moved

to the east, 23 reported that it moved to the west, 18 reported that it moved to the south,

and 4 reported that it moved to the north. (Thirty-six witness documents that reported a

streak of light did not contain information about the direction of travel.) 

Combined accounts (multiple witness reports summarized in a single document or

series of documents) presented another challenge. Although the Witness Group assigned

witness numbers to each witness identified in the documents, it was not always possible to

attribute specific observations to specific individual witnesses. The accounts of witnesses

who were interviewed in small groups, and especially married couples, were often

reported as if each person saw, remembered, and reported exactly the same events.

The Witness Group also determined that some documents included speculation

and conclusions about what witnesses observed without stating what the witness reported

observing. For example, one witness document stated, ìHe saw what he thought at the

451 It is also noteworthy that the documents recording the initial interviews of some witnesses, who have

since indicated that they believe a missile was used against the airplane, did not indicate that the witnesses

initially reported seeing a missile.
Factual Information 236 Aircraft Accident Report

time to be a large bright flare in the sky.î In this case, a description of exactly what the

witness observed was not provided, nor were other details, such as whether the flare was

moving, ascending, or descending, and if so, at what speed. The document only states

what the witness, located about 14.8-nm slant range from TWA flight 800, thought he was

observing. 

The Witness Group study report also noted that some witness documents contained

imprecise language. For example, one witness document stated that the witness reported

seeing fire falling from the sky and ìwhen it hit the water, approximately 8 to 10 seconds

later, there was a big explosion,î without clarifying whether the 8 to 10 seconds referred to

the amount of time between when the witness first saw the fire and when it hit the water or

between when the fire hit the water and the explosion. Another witness document stated

that the witness ìonly saw the object ascending for a split second when he observed an

orange or light colored object ascending from the southwest direction. He stated it

ascended for approximately five seconds,î without resolving whether the witness

observed one or two objects and whether it (or they) ascended for a split second or

5 seconds. In addition, documentation pertaining to some witnesses was internally

inconsistent. For example, one witness document indicated that the witness ìobserved

what appeared to be a USCG flareî and also that ìhe observed what he thought was a

shooting star.î Further, the document refers to the object as ìthe missile.î

The Witness Group study report determined that the witness documents indicated

that at least three witnesses characterized the origin of the streak of light differently over

two interviews. One witness document, based on a July 22, 1996, interview, stated that a

witness who was located about 22.3-nm slant range from TWA flight 800 observed a flare

rise straight up from the ocean. FBI agents reinterviewed this witness on June 26, 1997,

and the witness document based on that interview stated that the witness first observed the

flare about 15° to 20° above the horizon. Another witness (located about 14.7-nm slant

range from TWA flight 800) also characterized the origin of the streak of light differently

in two interviews. The witness document based on a July 23, 1996, FBI interview reported

that the witness saw a flare come from the water, whereas the witness document based on

an FBI interview conducted on July 15, 1997, reported that the witness saw a flare begin

ascending from 3 inches above the horizon. A third witness, interviewed by the FBI on

July 23, 1996, reported seeing ìwhat appeared to be cheap fireworks coming off the

beach.î The witness document based on a second FBI interview with this witness,

conducted on July 29, 1996, reported that the witness observed ìa white light that was

traveling skyward from the ground.î452 During subsequent interviews, this witness stated

that when he first observed the white light it was in the sky just above a house. 

452 During an April 30, 1999, briefing for the Witness Group, CIA analysts said that during a careful

reinterview by FBI agents, it became clear that the firework-like streak was in midflight when it was first

observed by this witness. The Safety Board received no documentation concerning this interview from the

FBI. 
Factual Information 237 Aircraft Accident Report

1.18.4.3 Research Regarding Accuracy of Witness Memory 

Research into human memory and its accuracy has demonstrated that human

memory is subject to error and that these errors occur for a variety of reasons.453 Further,

people tend to be unaware of their memory errors and may be overconfident in the

accuracy of their memories,454 and this confidence may increase over time.455 The Safety

Boardís own investigative experience has confirmed that witness reports are often

inconsistent with the known facts or with other witnessesí reports of the same events.456

1.18.4.3.1 Perception and Storing Information in Memory

People use their eyes, ears, and other sensory organs to take in information from

the environment; however, the raw data from the sensory organs are ìprocessed,î or acted

on, by the brain before people become aware of what their organs have sensed.457 The

brain receives and rapidly processes the sensory information such as light or sound that is

detected by the sensory organs.458 This processing occurs involuntarily459 and results in

the mental representation of sensation, which is called perception.460 In other words, the

brain forms perceptions from raw sensory data (for example, the brain perceives music or

words from sound energy). This processing is necessary because it filters and organizes

the many sights and sounds that people are exposed to into useful auditory and visual

images. However, because perception is not perfectly faithful to reality, people are subject

to illusions.461

453 Roediger, R. 1996. ìMemory Illusions.î Journal of Memory and Language. Vol. 35. p. 76-100.

454 Lindsay, R. C.; Wells, G. L.; and Rumpel, C. M. 1981. ìCan People Detect Eyewitness Identification

Accuracy Within and Across Situations?î Journal of Applied Psychology. Vol. 67. p. 79-89. 

455 Woocher, F. D. 1977. ìDid Your Eyes Deceive You? Expert Psychological Testimony on the

Unreliability of Eyewitness Identification.î Stanford Law Review. Vol. 29. p. 969-1,030.

456 For example, following the June 6, 1971, midair collision of a DC-9 with a Marine Corps F-4B, some

witnesses reported having seen the F-4B doing rolls and circling the area, although a surviving occupant of

the F-4B indicated that the only roll was performed at 15,500 feet, too far away to have been observed by

ground witnesses. (See NTSB/AAR-72/26.) Also, following the September 6, 1985, crash of Midwest

Express Airlines flight 105, a pilot witness reported seeing the airplane pitch up 20o, then down 12o, and

back up 20o, although the FDR data indicated no such pitch changes. (See NTSB/AAR-87/01.) In addition,

after the August 31, 1988, crash of Delta Airlines flight 1141 at Dallas Fort Worth International Airport, a

passenger on board the airplane stated that he saw the flaps extended, although the physical evidence from

the wreckage and airplane performance studies demonstrated, and the Board concluded, that the wing flaps

and slats had not been extended. (See NTSB/AAR-89/04.) Similarly, following the August 16, 1987, crash

of Northwest Airlines flight 255 at Detroit, Michigan, two pilot-rated witnesses reported seeing the flaps

extended, although the FDR indicated, and the Board concluded, that the flaps and slats were not extended.

(See NTSB/AAR-88/05.)

457 Coren, S. and Ward, L. M. 1994. Sensation and Perception. Harcourt Brace College Publishers. Fort

Worth, Texas.

458 Goldstein, E. B. 1989. Sensation and Perception. Wadsworth Publishing. Belmont, California.

459 Treisman, A. 1986. ìFeatures and Objects in Visual Processing.î Scientific American. Vol. 255.

p. 114-125.

460 Goldstein.

461 Coren and Ward.
Factual Information 238 Aircraft Accident Report

Perception of distance and motion

The visual image sensed by the eye is two dimensional, but the brain uses a variety

of cues to perceive depth and distance.462 These cues include information from binocular

disparity (the phenomenon that different views of the world are seen by each of a personís

two eyes) and the perception of perspective and texture gradients, which are present in any

feature-rich visual scene (such as looking down a busy city block).463 In the absence of

such depth-perception cues (for example, when looking across a desert or out over a large

body of water), people cannot judge distance accurately, and an object viewed under such

circumstances usually appears to be closer to the observer than it actually is.464 Further, in

the absence of depth perception cues, if structures, trees, or other objects are present near

an observersí line of sight to a critical object, people tend to perceive the critical object

and the intervening objects as equally distant from the observer.465 For example, a distant

tall building viewed along with several closer small buildings will appear to be located

about the same distance from the observer as the short buildings. This illusion has been

termed the equidistance tendency.466

Research also indicated that, if the critical object is in motion, the equidistance

tendency can also have the effect of exaggerating the perceived motion of the object.

Further, any motion of the observerís head may result in concurrent motion of the critical

object. Under viewing conditions that cause the critical object to appear closer to the

observer than it actually is, the critical object will appear to move in the same direction as

the observerís head moves.467 The closer the object appears to the observer, the greater the

amount of apparent motion.468 In other words, the perceptual error magnifies any

concurrent motion of the object caused by motion of the observerís head. Further, if the

critical object is actually in motion, the apparent motion that results from movement of the

observerís head may enhance, cancel out, or even reverse the motion of the critical

object.469

Storing information in memory

Researchers of human memory have recognized that in addition to memory for

specific events (episodic memory), human memory also contains memories that are not

462 Bruce, V. and Green, P. R. 1992. Visual Perception: Physiology, Psychology and Ecology. Lawrence

Erlbaum. London, England. 

463 Goldstein. 

464 Gogel, W. C. 1963. ìThe Visual Perception of Size and Distance.î Vision Research. Vol. 3.

p. 101-120.

465 Gogel, W. C. 1965. ìEquidistance Tendency and Its Consequences.î Psychological Bulletin. Vol. 64.

p. 153-163. 

466 Gogel. 

467 Gogel, W. C. and Tietz, J. D. 1973. ìAbsolute Motion Parallax and the Specific Distance Tendency.î

Perception and Psychophysics. Vol. 13. p. 284-302.

468 Gogel, W. C. and Tietz, J. D. 1977. ìEye Fixation and Attention as Modifiers of Perceived Distance.î

Perception and Motor Skills. Vol. 45. p. 343-362. 

469 Gogel, W. C. 1990. ìA Theory of Phenomenal Geometry and Its Applications.î Perception and

Psychophysics. Vol. 48. p. 105-123. 
Factual Information 239 Aircraft Accident Report

associated with specific events and places (semantic memory).470 These semantic

memories include generic information about the concepts that underlie everyday

activities. Humans use elements of semantic memory in an attempt to understand sensory

input. 471 This may occur without the observerís awareness.472 For example, someone

reading a newspaper does not carefully decipher the features of each letter to discover the

words that they form. Rather, the readerís knowledge of the subject matter and experience

with the language helps him or her anticipate the words and their meanings; this speeds

reading and allows it to become an almost automatic task.

Perception of events is driven by context cues and the observerís experience and

expectations,473 including semantic memory. This results in a fairly immediate

interpretation of the event. Thus, what is stored in memory about an event includes

fragments of the event itself and elements of interpretation.474

1.18.4.3.2 Retrieving Information from Memory

According to I. E. Hyman, a psychologist, ìPeople do not simply retrieve a

memory and replay the experienceómemory is not like a videotape. Instead people

combineÖknowledge from various sources with their own personal experiencesÖto

create memories.î475 Hyman has indicated that all memories are constructions. Research

has also shown that memories may correctly convey the gist of an event or a conversation;

however, some of the reported details are likely to be inaccurate and may reflect the

witnessís own beliefs or expectations about the event witnessed.476 Further, vivid and

confidently held memories of important events can change over time.477

Hyman has identified several types of memory errors, such as conjunction errors,

memory intrusions, time slice errors, and postevent information errors.478 Conjunction

470 Tulving, E. 1972. ìEpisodic and Semantic Memory.î Organization of Memory. Academic Press. New

York, New York.

471 Spiro, R. J. 1975. Inferential Reconstruction for Memory for Connected Discourse. Tech. Report

No. 2. Laboratory for Cognitive Studies in Education, University of Illinois. Urbana, Illinois.

472 Schneider, W. and Shiffrin, R. M. 1977. ìControlled and Automatic Human Information Processing I:

Detection, Search and Attention.î Psychological Review. Vol. 84. p. 1-66.

473 Lindsay, P. H. and Norman, D. A. 1972. Human Information Processing. Academic Press. New York,

New York.

474 Rumelhart, D. E., and Ortony, A. 1977. ìThe Representation of Knowledge in Memory.î Schooling

and the Acquisition of Knowledge. Lawrence Erlbaum Associates. Hillsdale, New Jersey.

475 Hyman, I. E. 1999. ìCreating False Autobiographical Memories: Why People Believe Their Memory

Errors.î Ecological Approaches to Cognition: Essays in Honor of Ulrich Neisser. Lawrence Erlbaum

Associates. Hillsdale, New Jersey.

476 Neisser, U. 1981. ìJohn Deanís Memory.î Cognition. Vol. 9. p. 1-22. Also, Hyman, I. E. and Neisser,

U. 1992. ìThe Role of the Self in Recollections of a Seminar.î Journal of Narrative and Life History. Vol. 2.

p. 81-103.

477 Neisser, U. and Harsch, N. 1992. ìPhantom Flashbacks: False Recollections of Hearing the News

about Challenger.î Affect and Accuracy in Recall: Studies of ìFlashbulbî Memories. Cambridge University

Press. New York, New York.

478 Hyman refers to postevent information errors as ìmisinformation errors.î
Factual Information 240 Aircraft Accident Report

errors occur when a witness blends information from different events into a single

memory. Memory intrusions (inclusion of specific information that was not originally

present) occur when people incorporate their general knowledge and experiences into

their memories and tend to occur in situations in which the recall of details is required or

expected.479 Time slice errors occur when people are asked to recall information about an

event, but they erroneously recall information about an event that took place at a different

time.480 In other words, they recall something that may be a true event, but something that

occurred before or after the event in question. Postevent information errors occur when

people receive some postevent information about an event that they have observed and

then later report the information as if they actually experienced it.481

Hyman has proposed three conditions that must exist for memory errors to occur:

witnesses must (1) believe that the event is plausible, (2) construct an image of the event,

and (3) believe that the event is a memory of something that they actually observed.482 A

plausible event is one that the witness believes could have actually occurred. For example,

research has demonstrated that people form false memories only when they consider them

to be plausible.483 As further discussed below, the more reliable or knowledgeable that the

witness believes the source of a suggested event is, the more likely the witness is to

believe that the event is plausible.

Hyman has indicated that plausible information only becomes a false memory if

witnesses construct a mental image of it. Other research supports the concept that

imagining an event increases the chances of witnesses erroneously claiming that they

recall it actually occurring.484 In fact, Hyman believes that almost any activity that causes

people to think about, imagine, or discuss events will lead to construction of images and,

thus, memories.

Finally, Hyman notes that for a false memory to occur, the witness must begin to

believe that the event is a memory of something that he or she actually observed. In other

words, the witness must misattribute the source of the incorrect information. Researchers

479 Hyman, I. E. and Rubin, D. C. 1990. ìMemorabeatlia: A Naturalistic Study of Long-term Memory.î

Memory and Cognitio. Vol. 18. p. 205-214.

480 Brewer, W. F. 1988. ìMemory for Randomly-selected Autobiographical Events.î Remembering

Reconsidered: Ecological and Traditional Approaches to the Study of Memory. Cambridge University Press.

New York, New York.

481 Belli, R. F. 1989. ìInfluences of Misleading Postevent Information: Misinformation Interference and

Acceptance.î Journal of Experimental Psychology: General. Vol. 118. p. 72-85; Loftus, E. F. 1979. ìThe

Malleability of Human Memory.î American Scientist. Vol. 67. p. 312-320; Loftus, E. F. 1979.

ìMentalmorphosis: Alterations in Memory Produced by the Bonding of New Information to Old.î Attention

and Performance IX. Lawrence Erlbaum Associates. Hillsdale, New Jersey. p. 417-434; and Loftus, E. L.

and Loftus, G. R. 1980. ìOn the Permanency of Stored Information in the Human Brain.î American

Psychologist. Vol. 35. p. 409-420.

482 Hyman. 1999.

483 Hyman, I. E. and Pentland, J. 1995. ìGuided Imagery and the Creation of False Childhood

Memories.î Journal of Memory and Language. Vol. 35. p. 101-117.

484 Goff, L. M. and Roediger, H. L. 1996. ìImagination Inflation: Multiple Imaginings Can Lead to False

Recollection of Oneís Actions.î Paper presented at the 37th annual meeting of the Psychonomic Society.

Chicago, Illinois.
Factual Information 241 Aircraft Accident Report

have demonstrated that people also often misattribute the source of information that they

have learned.485 That is, they may recall a particular piece of information, but not the

source of that information. For example, someone may read about an event in a newspaper

and also hear about it on television. Later, he or she may correctly recall details about the

event, but incorrectly recall from which source he or she learned the details. Researchers

have termed such events ìsource monitoring errors,î and research has shown that

eyewitness susceptibility to postevent information errors may result in part from source

monitoring errors.486 It has also been demonstrated that even when witnesses may be able

to accurately attribute information to a particular source (such as an interviewer or a

television news report), unless carefully questioned about the source of the information,

they may still report the information as if it had been personally observed.487

1.18.4.3.2.1 Postevent Information

Postevent information can be acquired in a variety of ways, including from

interviewers during official questioning, other witnesses, media information, and even

from the witnessís own thoughts as he or she ìreplaysî the event in memory. Research has

shown that an interviewerís phrasing of questions can have a considerable effect on the

information elicited from people.488 For example, the use of a definite article versus an

indefinite article in questioning (that is, ìDid you see the broken headlight?î versus ìDid

you see a broken headlight?î) is sufficient to increase the likelihood of a positive report.489

In another example, researchers found that peopleís estimates of speed were affected by

substituting different verbs (for example, ìcontacted,î ìbumped,î and ìsmashedî) for the

word ìhitî in the question ìHow fast were the cars going when they hit each other?î490

Researchers have also found that a witnessesí perception of the level of knowledge

or credibility of the interviewer affects their susceptibility to postevent information.

Witnesses who are asked a biased question by an interviewer whom they believe is

knowledgeable about the event are more likely to give an incorrect answer than (1)

witnesses who are asked a biased question by an interviewer whom they believe to be

naÔve or (2) witnesses who are asked unbiased questions.491 Smith and Ellsworth

conducted a study in which participants viewed a videotape of a simulated bank robbery

485 Johnson, M. K.; Hashtroudi, S.; and Lindsay, D. S. 1993. ìSource Monitoring.î Psychological

Bulletin. Vol. 114. p. 3-28. 

486 Lindsay, D. S. and Johnson, M. K. 1989. ìThe Eyewitness Suggestibility Effect and Memory for

Source.î Memory and Cognition. Vol. 17. p. 349-358. 

487 Lindsay and Johnson. 

488 Horgarth, R. M. (Ed.) 1982. Question Framing and Response Consistency. Jossey-Bass. San

Francisco, California.

489 Loftus, E. F., and Zanni, G. 1975. ìEyewitness Testimony: The Influence of the Wording of a

Question.î Bulletin of the Psychonomic Society. Vol. 5. p. 86-88.

490 Loftus, E. F. and Palmer, J. P. 1974. ìReconstruction of Automobile Destruction: An example of the

Interaction Between Language and Memory.î Journal of Verbal Learning and Verbal Behavior. Vol. 13. p.

585-589.

491 Smith, V. L. and Ellsworth, P. C. 1987. ìThe Social Psychology of Eyewitness Accuracy: Misleading

Questions and Communicator Expertise.î Journal of Applied Psychology. Vol. 72(2). p. 294-300.
Factual Information 242 Aircraft Accident Report

committed by more than one perpetrator.492 The participants who were asked, ìWhat did

the other guyís gun look like?î by a knowledgeable interviewer (one they who knew had

seen the videotape) in reference to an unarmed perpetrator were much more likely to

report a nonexistent gun than those participants who were asked the same question by a

naÔve interviewer (one who they believed had not seen the videotape), or those who were

asked, ìDid the other guy have a gun?î (regardless of the assumed/presumed knowledge

status of the interviewer).

In addition to information provided by interviewers, research has shown that

people will incorporate false information provided by other witnesses on a later recall test

of details about a written story.493 It has also been demonstrated that people may

incorporate details provided by others into their own descriptions of human faces.494

Research has also documented that, in a group interview about an event that the entire

group has observed, if misinformation is reported as a memory by a member of the group,

the likelihood is increased that others will also report this false information as if they had

also witnessed it.495 This research also showed that if an interviewer claims that other

witnesses to the same event (who are not present at the interview) have reported that a

given detail was part of the event, the witness being questioned is more likely to report

having seen the same detail, even though the detail was not part of the event.

The effect of subsequent repeated testing of the memories of subjects who were

exposed to misleading postevent information has also been studied. Research has

demonstrated that if witnesses reported the misinformation on the first test, they were

likely to repeat the misinformation on a second test.496 This, and other research497 appears

to show that repeated retrieval increases the probability of creating false memories.

Researchers have also found that likelihood of false memory creation is increased when

the postevent information is presented in multiple forms (reading, hearing, or seeing the

information).498

1.18.4.4 Witnesses of Special Interest

Because of their unique vantage points and/or the level of precision and detail

provided in the documentation pertaining to these witnesses, the following five witnesses

492 Smith and Ellsworth.

493 Betz, A. L.; Skowronski, J. J.; and Ostrum, T. M. 1996. ìShared Realities: Social Influence and

Stimulus Memories.î Social Cognition. Vol. 14. p. 113-140.

494 Loftus, E. F. and Green, E. 1980. ìWarning: Even Memory for Faces May be Contagious.î Law and

Human Behavior. Vol. 4. p. 323-334. 

495 Shaw, J. S.; Garven, S.; and Wood, J. M. 1997. ìCo-witness Information Can Have Immediate Effects

on Eyewitness Memory Reports.î Law and Human Behavior. Vol. 21. p. 503-523.

496 Roediger, H. L. et. al. 1996. ìMisinformation Effects in Recall: Creating False Memories Through

Repeated Retrieval.î Journal of Memory and Language. Vol. 35. p. 300-318.

497 Mitchell, K. J. and Zaragoza, M. S. 1996. ìRepeated Exposure to Suggestion and False Memory: The

Role of Contextual Variability.î Journal of Memory and Language. Vol. 35. p. 246-260. 

498 Mitchell and Zaragoza. 
Factual Information 243 Aircraft Accident Report

have generated special interest. Figure 41 is a map showing the debris zones and the

position of these five witnesses.

The pilot of Eastwind Airlines flight 507

According to radar data, about the time of the last secondary radar return from

TWA flight 800, Eastwind flight 507 was located about 24 nm northeast of TWA flight

800ís location at an altitude of about 15,400 feet.499 According to a statement that he

prepared on the night of the accident, the captain of Eastwind flight 507 reported seeing

what looked like an off-colored ìlanding lightî that at first he thought might be an engine

fire.500 He subsequently identified the ìlanding lightî as an airplane that he needed to

monitor. According to documents pertaining to an interview the day after the accident, the

captain of Eastwind flight 507 estimated that he observed the airplane for about 2 to

3 minutes.501 To signal the flight crew of the other aircraft that they were in sight, the

Eastwind captain switched on his own landing lights. As he flicked on his own landing

lights, the aircraft that he had been watching ìexploded into a very large ball of flames.î502

According to the statement that he prepared on the evening of the accident, ì[a]lmost

immediately two flaming objects, with flames trailing about 4,000 feet behind them, fell

out of the bottom of the ball of flame.î 

At 2031:50, the captain of Eastwind flight 507 reported to Boston ARTCC Sardi

Sector, ìwe just saw an explosion out here stinger bee five oh seven.î At 2037:20, he

further reported, ìah yes sir it just blew up in the air and then we saw two fireballs go

down to the to the waterÖalso ah there seemed to be a lightÖI thought it was a landing

lightÖand it was coming right at us at about I donít know about fifteen thousand feet or

something like that and I pushed on my landing lights ah you know so I saw him and then

it blew.î503 In the written statement that he made on the night of the accident, the captain

of Eastwind flight 507 stated that he immediately called ATC after seeing the ìexplosion.î

However, during an interview on March 25, 1999, he told Safety Board investigators that

after seeing the explosion, he paused for about 10 seconds before reporting the explosion

to ATC. 

In the statement that he made on the night of the accident, the captain of Eastwind

flight 507 also indicated that ìprior to the explosion I did not see any missile, air to air or

499 For each of the pilot witnesses mentioned in this section, the Safety Board used radar data to

determine positions and altitudes for witness aircraft relative to the last secondary radar return received from

TWA flight 800.

500 He indicated that he dismissed this possibility because he saw no smoke trail.

501 In a transcribed interview by the Safety Boardís Witness Group conducted about 3 years after the

accident, the witness indicated that he watched the light ìon and off for over two minutes, minimum but I

probably had seen him [the light] for over five minutes.î 

502 The captain indicated that the explosion occurred below his airplane at an altitude he estimated to be

between about 13,000 and 15, 000 feet msl. 

503 The captain of Eastwind flight 507 was the first airborne witness to report seeing the accident to the

air traffic controller at the Boston ARTCC.
Figure 41. A map showing the debris zones and the positions of the five witnesses of special interest.
Factual Information 245 Aircraft Accident Report

ground to air....î The pilot stated in a subsequent interview with the Safety Board that he

did not see anything coming up at or hitting the other airplane. 

The witnesses in the HH-60 helicopter

Crewmembers (the pilot, copilot, and flight engineer) in a New York Air National

Guard (NYANG) HH-60 helicopter, who at the time of the accident were performing

instrument approaches about 11 nm from where the main portion of the wreckage fell into

the ocean, were interviewed by the FBI shortly after the accident. According to

documentation of those interviews, the pilot of the NYANG helicopter stated that he first

observed a red streak of light moving very fast from right to left504 and that it was traveling

in a gradually descending arcóalmost horizontally. He stated that he observed the streak

for about 1 to 2 seconds after which he saw an explosion. He described the streak as

having the trajectory and image of a shooting star. In an interview with the Safety Boardís

initial Witness Group (conducted on January 11, 1997), he indicated that he observed a

red-orange streak of light moving horizontally for about 3 to 5 seconds in a gradually

descending arc that he described as a ìgentle descending trajectoryîósimilar to a

shooting star. He then saw what he described as an explosion, followed 1 to 2 seconds

later by a second and possibly a third explosion.505 He stated that he then saw a fireball,

which he described as ìfour times the size of the sun,î that hit the oceanís surface about

10 seconds after he first saw the fireball.506 The HH-60 pilot indicated that the fireball

followed the same trajectory as the streak. According to the witness documents, the pilot

stated that he had no idea what it was and that he never saw anything that he identified as

an airplane.

The witness documents indicated that the copilot and flight engineer from the

HH-60 stated that their first observation of the event was of flaming debris at an estimated

altitude of 4,000 to 5,000 feet and that this flaming debris hit the water about 8 seconds

after they first spotted it. According to the witness documents, the copilot stated that,

although he did not remember it on the evening of the accident, the next day he

remembered seeing an object streaking from his left to his right just before the appearance

of the fireball. He characterized this object as being like an ìincendiary deviceî or a

ìpyrotechnic.î He stated that he then saw a succession of three explosions, each longer

than the last.

The streak-of-light witness aboard USAirways flight 217507

Witness documents pertaining to interviews with a passenger seated in a window

seat on the right side of USAirways flight 217 (which was flying on a generally northerly

504 During his interview with the initial witness group, the pilot of the NYANG helicopter told

investigators that the helicopter was on a heading of 238o.

505 The witness described the explosions as being ìlike flakÖa hard explosion.î 

506 During questioning about a missile by TWAís representative to the Safety Boardís initial Witness

Group, the pilot stated, ìand when I looked up at this thing in the air, nothing said ëmissile.íî However, in

subsequent statements to the media, the pilot has stated that he saw an ìordinance explosionî and that he

believed that a missile was used against the airplane. 
Factual Information 246 Aircraft Accident Report

heading about 2.5 nm southwest of and about 8,000 feet above TWA flight 800 at the time

of the last secondary radar return) indicated that he ìobserved some blinking lights on a

small plane flying underneath the [USAirways flight 217] aircraftÖ[and] approximately

15-20 seconds later, he observed a light which appeared to be a ëflareíî for about

10 seconds moving from right to left. According to the witness documents, the witness

ìthen observed an initial small explosion in the same area that he last observed the flare.

Within a second later, the small explosion turned into a large explosionî and then a

descending fireball. 

Safety Board investigators used radar data to determine that the ìsmall planeî seen

passing underneath USAirways flight 217 by this witness was a U.S. Navy P-3 Orion.508

According to radar data and a visibility study done by Safety Board investigators, the

P-3 would have passed under USAirways flight 217 and out of view of the witness about

1 to 1 1/2 seconds before the end of the FDR data and the last radar transponder return

from TWA flight 800 (that is, about the time of the explosion in the CWT). The radar data

and visibility study also indicated that TWA flight 800 would not have been visible to

passengers seated on the right side of USAirways flight 217 when the CWT overpressure

event occurred.509

The witness on the Beach Lane Bridge in Westhampton Beach, New York

Witness documents pertaining to a witness who was standing on a bridge on Beach

Lane over Quantuck Canal in Westhampton Beach, New York (9.6-nm slant range from

the position of TWA flight 800ís last secondary radar return), stated that he saw ìa white

light that was traveling skyward from the ground at approximately a 40° angle.î He stated

that it looked like ìcheap fireworksî or a meteor that went up and ìëarched overí and

disappeared,î traveling in a south/southeasterly direction. However, in subsequent

interviews this witness apparently acknowledged that he did not see the light originate at

the surface of the earth because the light appeared in the sky above the house.510

The witness documents further indicated that the witness reported that he first saw

the light above the fourth or fifth house511 west of the bridge on the south side of Dune

Road and that the light was visible for about 15 seconds. He said that 2 or 3 seconds after

507 There were six additional witnesses aboard USAirways flight 217, five of whom were seated on the

right side of the airplane. (It could not be determined where the sixth witness was seated.) None of these

witnesses was classified by the Witness Group as having reported seeing a streak of light, but the documents

indicated that all of these witnesses reported seeing a large fireball on the right side of USAirways flight

217. The document pertaining to one of these witnesses (who was seated in seat 6F) stated that before the

witness saw a fireball, he saw an aircraft that ìhad what he took to be its landing lights on, and [he] could

observe cabin lights on the vessel.î According to the document, the witness reported watching the airplane

with ìcabin lightsî for 30 to 40 seconds, then seeing it bank to the left. The document then states, ìThe

aircraft exploded and a large round orange fireball appeared.î The document contains no mention of any

other light in the sky. 

508 The CIA reached the same conclusion.

509 For additional information regarding the radar data and visibility study, see Witnesses 4ñWitness

Visibility Study, dated June 22, 2000, and Witnesses 4ñViewing Angles Study, dated July 7, 2000. 

510 When asked how high above the house the light was, the witness indicated that, if one were to

imagine a flag pole on top of the house, the light would have appeared at the top of the flag pole.
Factual Information 247 Aircraft Accident Report

the white light disappeared he saw an orange fireball and ìthe fireball descended at

approximately a 30° angle and left a fire trail burning behind it.î The witness reported that

he saw the fireball descend behind the second house on the south side of Dune Road west

of the public parking lot and, seconds after first observing the fireball, he felt and heard an

explosion followed by three others. 

The Safety Board obtained aerial imagery of the bridge and the houses and used it

to draw sight lines. A sight line drawn from the witness position through the fourth or fifth

house intersected the flightpath of TWA flight 800 near the last transponder return from

the accident airplane. A sight line drawn from the witness position through the second

house intersected the green zone near where the aft fuselage was recovered. 

The witnesses on a boat near Great Gun Beach

Witness documents indicated that three witnesses were fishing in a boat about 1 to

1 1/2 miles512 from the shoreline (about 6 1/2 to 7 1/2 nm from the accident site). One of

those witnesses stated that he looked up and observed a faint yellow object, which he

thought might possibly be a meteor. This witness directed the attention of his companions

to the object. He stated that it looked like a ìshooting starî moving in a downward,

45° sloping arc and that it that grew larger and brighter. He stated that he then heard a

ìrushing roarî type of sound and at that moment realized that he was watching an airplane

and that he saw one wing separate from the fuselage and remaining wing513 at the same

time as he heard and felt514 an explosion. The witness reported, ìIt [the airplane] had

rolled over onto its left side, and seemed to be hooking northward and downward.î He

stated that he then saw what he believed was the right wing separate from the fuselage.

The other two witnesses in the boat reported that, before the object broke apart, it seemed

to turn toward the north. One of the other two witnesses estimated that he observed part of

the object separate from the remainder about 8 seconds after he first sighted the light. (The

witness who saw the ìshooting starî estimated that he began observing it before his

companions saw it.)

1.18.4.5 Witness Line-of-sight Study 

A line-of-sight study was conducted by the Suffolk County Police Departmentís

Marine Bureau to determine the geographical area from which witnesses reported seeing a

flarelike object.515 The study was accomplished between July 20 and 30, 1996, during FBI

reinterviews of several witnesses who reported seeing a streak of light or flarelike object

511 The witness document pertaining to a July 23, 1996, FBI interview states that the witness stated that

he ìsaw what appeared to be cheap fireworks coming off the beach about 4 or 5 houses west of the bridge.î

A document pertaining to a more detailed FBI interview on July 29, 1996, states that the ìwhite light came

in to view just above the roof top of the fourth house west of the public parking area on Dune Road.î

512 In one document of an interview with the witness who first saw the object, the witness estimated that

he was about 1 1/2 miles from the shoreline. In a second document, the witness stated that he was about

1 mile offshore. 

513 The witness stated that he saw the right wing separate from the fuselage; however, the Sequencing

Group determined that the left wing separated from the fuselage and right wing. 

514 He reported that he felt a concussion on his chest at the same time as he saw the wing separate.
Factual Information 248 Aircraft Accident Report

on the evening of July 17, 1996. During these reinterviews, the witnesses were taken to

the locations from which they made their observations. The position (latitude and

longitude) of each witness and the horizontal direction (azimuth) were recorded for each

witness location to the area where he or she observed the streak of light or flarelike

object.516 Data from witnesses who could describe their positions and their observations of

the flarelike object relative to fixed reference points (such as a flagpole or a building) were

recorded and then plotted on a map of the accident debris fields. Lines depicting the

azimuth from the witnessís position to the area that the witness reported seeing the streak

of light or flarelike object were also drawn. 

The line-of-sight study determined the latitude and longitude of two points at

which the sight lines appeared to intersect (according to the Suffolk County Police

Departmentís Marine Bureau) and recommended that the FBI search a 1-nm area around

these points for evidence of a missile.517 Safety Board investigators also plotted these data,

developed a sight line plot map using computer map-making software, and determined

that 7 of the 11 sight lines intersected the green zone debris field and that two passed just

north of it. Figure 42 shows a map of the sight lines plotted for 11 of the witnesses.

1.18.4.6 Central Intelligence Agency Review of the Witness Data

Early in the investigation, at the FBIís request, the CIA reviewed data regarding

the TWA flight 800 accident (including witness reports and information; FDR, CVR,

ATC, and radar data; and infrared sensor information from a U.S. satellite) to determine if

the streak of light reported by many of the witnesses might have been a missile. The CIAís

video518 stated that ìnot surprisingly, most eyewitnesses saw only the most conspicuous

segment of the disasteróthe fire trail and cascade of flames in the 10 to 15 seconds before

the aircraft hit the water.î 

The CIA video indicated that because ìsound from the aircraftís explosion took

more than 40 seconds to reach each of the 58 eyewitnesses [that they evaluated519] who

reported hearing sounds associated with the disaster. Therefore, any events eyewitnesses

reported seeing at about the time they heard the first sound took place well after the

explosion.î The CIA video concluded that ìmost [witnesses] describe observations made

only during the final moments of the disasterówell after the aircraft explodedÖ.To date,

there is no evidence that anyone saw a missile shoot down TWA flight 800.î 

515 The FBI requested this study as part of its investigation of the possibility that a missile had been used

against the accident airplane.

516 A Garmin Model 45 GPS receiver was used to determine position (in degrees, minutes, and seconds),

and a hand-held magnetic compass was used to determine azimuth (in magnetic degrees).

517 These two locations were well within the undersea area covered by the 9-month search and recovery

operation that concluded in April/May 1997. (See appendix C for information regarding the search and

recovery operation.) 

518 For additional information, see the CIA video, ìTWA Flight 800: What did the witnesses see?,î first

shown at the November 18, 1997, FBI press conference. 

519 The FBI provided the CIA with only some of the witness documents that were later provided to the

Safety Board.
Figure 42. Map of sight lines plotted for 11 witnesses. 
Factual Information 250 Aircraft Accident Report

1.18.4.7 Missile-Related Information

At the Safety Boardís request, the Naval Air Warfare CenteróWeapons Division

(NAWC-WD) at China Lake, California, provided assistance in identifying possible

missile impact damage on the accident airplane wreckage. NAWC-WD experts conducted

detailed examinations of the recovered wreckage and reconstructions and analyzed other

data to evaluate the possibility that a shoulder-launched missile destroyed the airplane.520

According to a ballistics/warhead testing expert from the NAWC-WD, if a

missile521 with a live warhead impacted the airplane and detonated, the wreckage would

likely have exhibited extensive damage to the impact area from the initial penetration of

the missile and the subsequent dispersion of high-velocity postdetonation fragments. He

described the following four regions522 of warhead damage as being typical for nearly all

missile warheads: 

ï The first region, which is in the immediate vicinity of the detonation, is

characterized by complete removal of the structureÖdue to weakening by

fragment penetrations and dislocation by the blast pressure wave. 

ï The second region contains numerous high-velocity[523] fragment penetrations,

soot residue from the explosive, and may exhibit distortion, but not widespread

removal, of structureÖdue to the pressure wave. 

ï The third region contains more widely spaced high-velocity fragment

penetrations, usually with little evidence of sooting or pressure wave damage.

ï The fourth region contains only occasional high- and low-velocity fragment

penetrations. 

520 For detailed information, see the NAWC-WD report, ìTWA Flight 800 Missile Impact Analysis,î

dated December 1997. 

521 The expert considered it most likely that, if a missile were involved in the TWA flight 800 accident, it

would be a shoulder-launched missile (a Man Portables Air Defense System) because (1) shoulder-launched

missiles are available and are effective terrorist weapons and (2) larger surface-to-air and air-to-air missile

systems are expensive, require training and equipment not readily available, and leave clearly identifiable

evidence over larger areas of the target than shoulder-launched systems. 

522 According to the NAWC-WD expert, these regions did not have distinct boundaries between each

other, and overlapping of characteristic damage between regions is common. 

523 The expert defined high-velocity as ìin excess of around 4,000 feet per secondî and indicated that

holes resulting from high-velocity penetrations would be ìvisually quite different from low-velocity

impacts.î 
Factual Information 251 Aircraft Accident Report

According to the NAWC-WD ballistics/warhead testing expert, although some

types of damage inflicted by warheads can be caused by other events524

high-velocity fragment penetrations are unique to explosive events and give

investigators a conclusive method of identifying these encounters when they

occur. Previous testing on both commercial and military aircraft has shown that

even with small shoulder-launched weapons, high-velocity fragmentation damage

to the aircraft will exist over large areas of the targetÖ.Based on previous testing

performed by the military and [FAA], it is inconceivable that a warhead could

have detonated in or near the fuselage without leaving evidence of high-velocity

fragmentation damage somewhere on the recovered wreckage. 

NAWC-WD also performed infrared signature measurements of 747s on approach

to a commercial airport to identify the most likely tracking points for a shoulder-launched

missile. According to the NAWC-WD report, infrared seekers on shoulder-launched

missiles typically aim for hot spots, such as engine exhausts or inlets. The NAWC-WD

infrared measurements showed that the four engines and the underside of the fuselage

where the air conditioning pack bay is located (under the CWT) were the ìhot spotsî on

the 747. The NAWC-WD report indicated that its experts observed no evidence of

high-velocity fragmentation damage on the accident airplane. Specifically, the report

indicated that no such damage was observed on the accident airplaneís engines or air

conditioning pack components. 

The NAWC-WD report concluded the following:

No conclusive evidence of missile impacts exists on any of the recovered

wreckage of TWA flight 800. No evidence of high-velocity fragment impacts

exists, which indicates a live warhead did not detonate within or near the exterior

of the aircraft....The possibility that a shoulder-launched missile was launched at

TWA flight 800, failed to intercept it, self-destructed in close proximity,[525]

 and

initiated the breakup of the aircraft is highly improbable.î 

The Safety Board used missile performance data provided by NAWC-WD,

warhead detonation data from the Defense Intelligence Agencyís Missile and Space

Intelligence Center (MSIC), and damage tolerance data from the

Survivability/Vulnerability Information Analysis Center (SURVIAC) in a computer

simulation study to further evaluate the possibility that a shoulder-launched missile aimed

at the accident airplane did not reach its target but selfdestructed at a point close enough

for a fragment to enter the CWT.526

According to the Safety Boardís study, for missile selfdestruction to be a possible

source of CWT ignition for TWA flight 800, the missile must be (1) close enough at

detonation for a fragment to enter the CWT, (2) oriented such that the fragmentation

524 For example, soot deposits may be caused by a fuel-fed fire, and low-velocity penetrations may be

caused by ground/water impacts. 

525 According to missile warhead experts, shoulder-launched missiles will selfdestruct their warheads

about 15 to 17 seconds after they are launched if they do not reach a target. 

526 For additional information, see Missile Self Destruct Performance Study, dated April 13, 2000. 
Factual Information 252 Aircraft Accident Report

pattern intersects the CWT, and (3) far enough away at detonation that the airplane was

not peppered with holes.527

In an effort to bracket the possible missile performance range, NAWC-WD

provided the Safety Board with data for a low-performance and a high-performance

missile.528 According to NAWC-WD personnel, a typical missile flight would consist of

about 8 seconds of engine burn (7.8 and 8.4 seconds for the two missiles the Board

received data for) followed by a high-speed coast of about 7 to 9 seconds to

selfdestruction if no target has been hit. During the 15- to 17-second flight, the missile

would track to a heat source on the airplane. 

According to the MSIC, warhead detonation typically results in about

1,000 fragments of various sizes, with an average mass of less than 10 grains

(0.0229 ounce), and a mass as large as 40 grains (0.0916 ounce) in the pattern shown in

figure 43. MSIC reported that warhead fragments could have an initial velocity of up to

9,000 fps. Data obtained from the SURVIAC indicated that 10- and 40-grain fragments

would require velocities of 3,500 fps and 2,000 fps, respectively, to penetrate the CWT if

they impacted the bottom of the CWT at a right angle. Using the initial speed and speed

decay rate for the warhead fragments, the Safety Board calculated that a detonation would

have to occur 40 feet or less below the airplaneís CWT for a fragment to retain sufficient

velocity to penetrate the CWT.

SURVIACís data indicated that an even greater velocity would be required for a

fragment to penetrate the CWT if it struck the bottom of the CWT at an angle. For

example, a 40-grain fragment impacting the CWT at a 60° angle would require a velocity

of more than 4,500 fps to penetrate the bottom of the tank. Using the initial speed and

speed decay rate for the warhead fragments, the Safety Board calculated that a detonation

would have to occur within 17 feet of the airplaneís CWT for a fragment to retain

sufficient velocity to penetrate the CWT at a 60° angle. Further, SURVIAC data showed

that a fragment loses both velocity and mass when it must pass through other layers of

material (such as metal or fiberglass) before penetrating the component of interest (in this

case, the CWT). Additionally, when the fragment speed decay rate was adjusted to

represent a fragment passing through Jet A fuel, calculations indicated that the fragment

would slow from 9,000 fps to 2,000 fps in less than 1 inch. The Board notes that,

regardless of where the detonation occurred, a missile fragment would have had to have

penetrated multiple layers of material to enter the CWT. 

527 As previously discussed, examination of the wreckage revealed only two holes in the horizontal

pressure deck (above the left side of the wing landing gear bay) that contained some features of both

high-velocity and low-velocity penetrations. However, the features of these penetrations indicated that they

were made by objects coming from different directions; therefore, it is not likely that they originated from a

common location. Moreover, the physical features of both holes indicated they were made by objects

traveling from inside the airplane out, rather than the reverse. 

528 Because of the sensitive nature of the data, the missiles for which data were provided were not

identified. 
Factual Information 253 Aircraft Accident Report

The Safety Boardís computer simulations (using this missile performance data and

taking into account the winds at the time of the accident) showed that the locations on the

surface from which a missile launched at TWA flight 800 would result in selfdestruction

within 40 feet of the airplane were extremely limited. Figure 44 shows airplane locations

at which, if a high-performance missile were launched at the airplane, the missile would

selfdestruct within 40 feet of the airplane. Note that, at the scale shown, the launch area

that would result in a selfdestruction within 40 feet of the airplane is not as wide as the line

showing the 15-second perimeter in this figure.

Simulations further showed that, because of the pattern in which missile fragments

are dispersed after selfdestruction (see figure 43), launches from many of these positions

would result in no fragments having a trajectory that would result in penetration of the

CWT. For example, because of the nature of missile guidance, a launch from most

positions on the perimeter shown in figure 44 would result in the missile catching the

aircraft from behind. The area in front of the missile is not in the fragment dispersal area. 

Figure 43. Diagram of warhead detonation fragment pattern.
Factual Information 254 Aircraft Accident Report

Figure 44. Diagram showing airplane and missile launch site locations that could result 

in missile selfdestruct within 40 feet of the plane.

1.18.4.7.1 Missile Visibility Study

On the basis of performance data for a fairly high-performance, shoulder-launched

missile, to collide with TWA flight 800, a hypothetical missile would have had to have

been launched between about 4.5 nm and 11 nm from shore. Safety Board investigators

consulted with missile experts in U.S. intelligence and defense communities to determine

if the launch and flight signatures of a shoulder-fired missile would be detectable at those

distances. Weapons systems experts indicated that there were no data available on which

to base an answer to this question. Consequently, the Board conducted (and videotaped)

missile visibility tests on April 28, 2000, at Eglin Air Force Base, Fort Walton Beach,

Florida, to determine if observers could see and describe the characteristics of a missile

launch.

The missiles were launched just after sunset on April 28, 2000, when the surface

weather conditions and the position of the sun were similar to the conditions on Long

Island at the time of the accident. Observer groups were positioned every 2 nm from the

launch point out to 14 nm.529 The observers were told where to look, and they listened to

the launch countdowns via handheld radios. Although nine countdowns were

TWA flight 800 position at launch for encounter at 15 sec

Launch site

TWA flight 800

Sam2 Missile Footprint
Factual Information 255 Aircraft Accident Report

accomplished, only three missiles were launched. The observers did not know which

countdowns would lead to launches and which would not. Statements and other

information were obtained from the observers between countdowns. 

The three missile launches were easily detected by all of the observers. There were

four false alarms (reports of something moving in the sky after a countdown that was not

followed by a missile launch). According to the observers, the launched missiles appeared

as a bright light or flame rising from the surface. The observers also indicated that the light

was only visible during the (about 7 to 8 second) burn, and then they observed nothing

until the missile selfdestructed about 7 to 8 seconds later.530 Observers who reported

seeing the missile selfdestruct described it as an abrupt flash of light. Some of the closest

observers reported hearing a muffled boom about 20 to 30 seconds after they observed the

flash of light. All observers reported that they were unable to see the body of the missile

from any range.531

529 The observers who were positioned 14 nm from the launch site were originally positioned closer to

the missile launch site (12 nm out), with instructions to move farther out for subsequent launches if the

missile was visible from the closer position during the first observed launch. Thus, the lighting conditions

for the tests during which these witnesses were positioned 14 nm from the launch site were darker than in

earlier tests. For the first launch, which was conducted under lighting conditions that were comparable to

those that existed at the time of the TWA flight 800 accident, no observers were positioned farther than

12 nm from the launch site. 

530 Not all observers reported observing the missile selfdestruct. For example, during the first launch,

five of the six observers positioned within 6 nm of the launch site reported observing the missile

selfdestruct. None of the observers positioned more than 6 nm from the launch site reported observing the

missile selfdestruct. 

531 For the second launch, one observer at 10 nm replied, ìmaybe,î when asked if he saw the missile

body. He answered negatively for the other two launches, including the first launch, which was conducted

under lighting conditions comparable to those that existed at the time of the TWA flight 800 accident.
256 Aircraft Accident Report

2. Analysis

2.1 General

The flight crew was properly certificated and qualified and had received the

training and off-duty time prescribed by Federal regulations. No evidence indicated any

preexisting medical or behavioral conditions that might have adversely affected the flight

crewís performance during the accident flight.

The airplane was certificated, equipped, and dispatched in accordance with Federal

regulations and approved TWA procedures. (Maintenance issues relating to airplane

wiring are discussed in section 2.5.)

At the time of the accident, there were light winds and scattered clouds in the area,

but there were no significant meteorological conditions that might have disrupted the

flight. 

2.2 The Accident Sequence

2.2.1 The In-Flight Breakup

On July 17, 1996, about 2019, TWA flight 800 took off from John F. Kennedy

International Airport (JFK), New York, New York, destined for Charles DeGaulle

International Airport, Paris, France. The airplane climbed over the Atlantic Ocean without

incident and leveled off at its assigned altitude of 13,000 feet about 2027. At 2030:18, in

response to an air traffic control (ATC) instruction to climb to 15,000 feet, the flight crew

increased the airplaneís engine thrust. The airplane was ascending through 13,760 feet at

2031:12 when both the cockpit voice recorder (CVR) and the flight data recorder ceased

recording without warning.

Many witnesses in the area of the accident at the time that it occurred reported that

they saw and/or heard an explosion, accompanied by a fireball over the ocean, and

observed debris falling to the water. These witness reports and the widespread distribution

of the wreckage indicated that TWA flight 800 had experienced a catastrophic in-flight

structural breakup. In addition, a noise recorded on the CVR in the last few tenths of a

second before the CVR recording stopped was similar to the last noises heard on CVR

recordings from other airplanes that had experienced structural breakups (including fuel

tank explosions).532 On the basis of this initial information, investigators considered

several possible causes for TWA flight 800ís in-flight structural breakup: a structural

failure and decompression; detonation of a high-energy explosive device, such as a bomb

532 For more information about CVR-related tests, see section 1.11.1.2.
Analysis 257 Aircraft Accident Report

exploding inside the airplane or a missile warhead exploding upon impact with the

airplane; and a fuel/air explosion in the center wing fuel tank (CWT). 

2.2.1.1 Consideration of a Structural Failure and Decompression

Close examination of the wreckage revealed no evidence of preexisting airplane

structural faults (such as fatigue, corrosion, or mechanical damage) that could have

contributed to the in-flight breakup. The examination revealed that the structure did have

minimal preexisting corrosion damage, none of which could have led to or affected the

breakup of the airplane. Small fatigue cracks were found in some parts of the airplane,

including in the lower chord of the front spar and in the shear ties for the floor beams and

stiffeners at the front spar; however, none of these cracks had coalesced into a propagating

crack that could have led to the in-flight breakup. Further, although the joint between

fuselage sections 41 and 42 on some 747s purportedly had been subject to manufacturing

assembly problems, there was no evidence that it had separated in any locations before

impact.

It was also suggested that the breakup could have been initiated by the in-flight

separation of the forward cargo door. However, all eight of the latching cams along the

bottom of the door (and some pieces of the cargo door itself) remained attached to the pins

along the lower door sill, and there were no indications of preimpact failure of the hinge at

the top of the door. This evidence indicates that the door was closed and locked at impact.

Further, deformation and fracture patterns on the door matched damage to the adjacent

fuselage structure, confirming that the door was in the closed position at the time of

impact. Therefore, the Safety Board concludes that the in-flight breakup of TWA flight

800 was not initiated by a preexisting condition resulting in a structural failure and

decompression. 

2.2.1.2 Consideration of a High-Energy Explosive Device Detonation (Bomb 

or Missile Warhead)

Several factors led to speculation that the accident might have been caused by a

bomb or missile strike. These factors included heightened safety and security concerns

because of the 1996 Olympics then being held in the United States, the fact that TWA

flight 800 was an international flight, and the sudden and catastrophic nature of the

in-flight breakup. In addition, numerous witnesses to the accident reported seeing a streak

of light and then a fireball, which some people believed represented a missile destroying

the airplane.533 Further, some anomalous primary radar targets were recorded by the Islip,

New York, radar site in the general vicinity of TWA flight 800 at the time of the accident

that apparently could not be explained.534 Accordingly, the Safety Board considered the

possibility that a bomb535 exploded inside the airplane or that a missile warhead from a

shoulder-launched missile exploded upon impact with the airplane.536

Testing performed by the Federal Bureau of Investigation (FBI) found trace

amounts of explosives on three separate pieces of airplane wreckage (described by the FBI

533 For more information about these witness reports, see section 2.2.2.
Analysis 258 Aircraft Accident Report

as a piece of canvaslike material and two pieces of floor panel). However, none of the

damage characteristics typically associated with a high-energy explosion of a bomb or

missile warhead (such as severe pitting, cratering, petalling, or hot gas washing) were

found on any portion of the recovered airplane structure, including the pieces on which the

trace amounts of explosives were found. Only about 5 percent of the airplaneís fuselage

was not recovered, and none of the areas of missing fuselage were large enough to have

encompassed all of the damage that would have been caused by the detonation of a bomb

or missile. Although several large holes are visible in the reconstructed portion of the

airplane fuselage, almost all of the structure that originally filled in these holes is attached

to the remaining structure but is folded either inward or outward. No area of structure in

the reconstructed portion of the airplane contained any unexplained holes large enough to

represent the entry point of a missile.537 Further, the victimsí remains showed no evidence

of injuries that could have been caused by high-energy explosives,538 nor was there any

damage to the airplane seats and other interior components consistent with a high-energy

explosion. 

Investigators considered several scenarios to determine how the trace amounts of

explosive residue might have gotten on the wreckage from the accident airplane. Trace

amounts of explosive residue could have been transferred to the contaminated pieces from

the military personnel (and their associated clothing, boots, and equipment) that were on

534 Investigators closely evaluated three sequences of primary radar returns recorded by the Islip radar

site around the time of the accident (from about 1 minute before the accident to 9 minutes after), which

appear to show three targets moving at 300 to 400 knots ground speed about 10 nautical miles (nm)

southwest of the accident. None of the three sequences intersected TWA flight 800ís position at any time,

and all of them were moving away from the accident airplane. For each of these three sequences of primary

returns, investigators found that the signal strengths varied randomly from very high to very low, the

azimuth was always 150o to 160o relative to the radar, and there were no primary or secondary tracks leading

to or away from them. Further, Islip was the only radar site out of six primary radar sites covering this

airspace that recorded these primary returns. Investigators learned from air traffic controllers and radar

technicians that ground or building reflections of primary radar returns from aircraft flying in one

geographic area can cause ìfalse primaryî targets to be recorded as though the aircraft were flying in another

geographic area. Numerous buildings and structures in the area around the Islip radar site could have created

such reflections. Investigators identified commercial airplanes traveling through other areas within the

coverage of the Islip radar at the same time that the three sequences of primary returns were recorded; these

airplanes had similar ground speeds and flightpaths as the three sequences of primary radar returns. These

findings indicate that the three sequences of 300 to 400 knot primary radar returns recorded by the Islip radar

site between about 2030 and 2040 do not represent unexplained objects (such as a missile) but, rather,

represent false or reflected returns from actual aircraft in other geographic areas.

535 For the purposes of this discussion, the term ìbombî means an explosive device designed to release

destructive material at high velocity upon detonation. The term does not refer to an explosive device

designed only to set off a small charge of sufficient strength to penetrate the fuel tank and ignite explosive

vapors. The possibility that such a device might have ignited the fuel/air vapor in the CWT is discussed in

section 2.2.1.2.

536 The possibility that the CWT was ignited by a missile fragment from a missile warhead that exploded

at a sufficient distance from the airplane such that most of the other fragments would not have caused the

type of damage typically left by an exploding missile warhead is discussed in section 2.3.1.3.

537 The largest missing area in the portion of the airplane fuselage recovered from the red zone (where the

first pieces to depart the airplane would have landed) was about 1 square foot. A missile warhead is about

3 inches in diameter; however, the underlying structural pieces (stringers and/or frames) in each of the areas

of missing fuselage showed no signs of penetration by a missile warhead.
Analysis 259 Aircraft Accident Report

board the accident airplane when it was used to transport troops during the Gulf War in

1991. In addition, explosives were placed and then removed from several locations in the

accident airplane during a dog-training explosive detection exercise about 1 month before

the accident.539

However, testing by the Federal Aviation Administrationís (FAA) Technical

Center indicated that residues of explosives found on the accident airplane would dissipate

completely after 2 days of immersion in sea water.540 Very few pieces of airplane

wreckage were recovered during the first 2 days after the accident, and those pieces that

were recovered were found floating on the oceanís surface; several days of side-scan sonar

searching preceded any underwater wreckage recovery operations. Therefore, it is very

likely that the pieces on which the explosive residues were found were immersed in ocean

water for considerably more than 2 days before they were recovered. 

Thus, it is quite possible that the trace amounts of explosive residue detected on

wreckage pieces from the accident airplane were not present before the airplane entered

the water but, rather, were deposited during or after recovery operations. The military

personnel, ships, and ground vehicles used during the recovery operations had come into

frequent previous contact with explosives. Trace amounts of those substances could have

been transferred from the surfaces of the ships or ground vehicles, or from clothing and

boots of military personnel, onto wreckage pieces during the recovery operations or

through subsequent contact with the pieces in the airplane hangar where the airplane

wreckage was later assembled and laid out. 

Despite being unable to determine the exact source of the trace amounts of

explosive residue found on the wreckage, the lack of any corroborating evidence

associated with a high-energy explosion indicates that these trace amounts did not result

from the detonation of a high-energy explosive device on TWA flight 800. Accordingly,

the Safety Board concludes that the in-flight breakup of TWA flight 800 was not initiated

by a bomb or a missile strike. 

2.2.1.3 Consideration of a Fuel/Air Explosion in the Center Wing Fuel Tank

It was clear from the wreckage recovery locations that the first pieces to depart the

airplane were from the area in and around the airplaneís wing center section (WCS),

538 According to a medical/forensic analysis report prepared for the Department of Justice/FBI by a

medical/forensic expert, all injuries were consistent with a severe in-flight breakup and subsequent water

impact.

539 During that exercise, an airport police officer placed five training aids containing explosives (one of

which was cracked and could have released small amounts of explosives) in various locations inside the

passenger cabin. The officerís activities were limited to the forward portion (forward of row 30) of the

airplane. Training aids were placed in the following five locations: (1) in row 1, in the center armrest

compartment separating seats 1 and 2; (2) in a small closet at the rear of the upper deck; (3) in row 10, inside

the rear pouch of the seatback in front of seat 9; (4) in row 20, in the overhead compartment above seats 1

and 2; and (5) beside exit door R3.

540 The testing included both cloth and aluminum airplane pieces that were contaminated by manually

deposited explosives and by exposure to an actual explosion. In both cases, no trace of explosives remained

after 2 days of immersion in sea water.
Analysis 260 Aircraft Accident Report

which includes the CWT, and, therefore, that the breakup must have initiated in this area.

Specifically, wreckage found in the red zone (the wreckage zone closest to JFK along the

airplaneís flightpath and, therefore, containing the earliest pieces to depart the airplane)

consisted primarily of pieces from the WCS front spar and spanwise beam (SWB) 3, the

manufacturing access door from SWB2,541 the two forward air conditioning packs, large

pieces of a ring of fuselage structure just in front of the wing front spar, and main cabin

floor beams and flooring material from above the WCS and from the fuselage in front of

the WCS. None of these pieces showed evidence of prolonged fire exposure. (There was

insufficient fuel in the CWT for its destruction to have resulted in a significant fire.)542

However, some pieces were lightly sooted, indicating that there was some fire associated

with the initial event in this area.

As discussed in section 1.16.3, the Metallurgy Structures/Sequencing Groupís

sequencing study concluded (in part because pieces from inside the CWT were among the

first pieces to depart the airplane) that the initial event in the breakup sequence was an

overpressure event543 within the CWT and that the earliest piece of the airplane to be

disturbed was SWB3. Specifically, the Sequencing Group concluded that SWB3 fractured

at its upper end and that overpressure within the CWT caused it to rotate forward about its

lower end. As the upper end of SWB3 rotated forward, it impacted the aft surface of the

front spar, leaving distinct witness marks across most of the front spar. Analysis of the

wreckage further indicated that when SWB3 impacted the front spar, it initiated multiple

fractures along the upper chord of the front spar. Overpressure escaping from the CWT

caused the front spar to bulge forward on either side of the two potable water bottles

attached next to each other at the center of the front spar. The upper end of the front spar

then completely separated from the upper skin of the WCS. After the upper end of the

front spar was completely separated from the upper skin, the overpressure remaining

within the CWT forced the WCS lower skin and the forward end of the keel beam

downward. Downward loading of the forward end of the keel beam greatly increased the

stress in the ring chord and in the fuselage skin adjacent to the front spar.

As the keel beam was forced downward, cracking propagated down through the

lower pressure bulkhead and ring chord and immediately entered the fuselage skin at

stringer 40 right. This initial fuselage skin crack at stringer 40 right propagated forward at

an angle to the left (toward the bottom center of the airplane), then branched

circumferentially left and right, and then (from the left branch) propagated back toward

the front spar on the lower left side. Thus, at this point, the fuselage skin had cracked

541 The remainder of SWB2 was found in the green zone (the wreckage zone farthest from JFK along the

airplaneís flightpath, and, therefore, containing the last pieces to depart the airplane). The manufacturing

door was only lightly sooted, whereas the sooting on the remainder of SWB2 was moderate to heavy,

indicating that the remaining portion of the WCS was subjected to a more substantial fire after the early

departure of the manufacturing door.

542 Therefore, the large ìfireballî described by some witnesses could not have been caused by the initial

destruction of the CWT. For additional discussion about the nature of the reported fireball, see

section 2.2.2.1.

543 As previously mentioned, the Sequencing Group defined an overpressure event as ìone that would

create sufficient pressure in a relatively short time to a level at which the structural integrity of the CWT is

compromised.î
Analysis 261 Aircraft Accident Report

around three sides of a large piece of fuselage structure (including the piece identified

after recovery operations as ìLF6Aî); only the side adjacent to the lower pressure

bulkhead (beneath the front spar) remained attached to the other structure. 

Downward loading (from normal cabin pressurization and vented CWT

overpressure) on this weakened fuselage piece was transmitted to the forward end of the

keel beam, which caused it to separate from the WCS lower skin panel and fracture about

22 inches aft of the mid spar. The separation of the large piece of fuselage created a large

opening in the fuselage, just forward of the front spar, through which pieces of the front

spar, SWB3, and the manufacturing access door from SWB2 exited the airplane. (The

apparent forward and upward motion of the manufacturing access door after it separated

from SWB2 at its lower edge indicates that, at that time, the pressure on the aft surface of

the door was greater than on the forward surface, consistent with venting of the

overpressure forward of SWB2 through the large opening in the bottom of the fuselage.)

The sequencing study determined that the breakup sequence was initiated by an

overpressure inside the CWT. Because there was no evidence that a high-energy explosive

device detonated in this (or any other) area of the airplane, this overpressure could only

have been caused by a fuel/air explosion in the CWT.

Safety Board flight tests that recreated, to the extent possible, the conditions

experienced by the accident airplane indicated that fuel vapor temperatures within the

CWT at the time of the accident ranged from 101° to 127° F. Further, Safety

Board-sponsored testing at the CIT showed that Jet A fuel vapors under conditions

simulating the pressure, altitude, and fuel mass loading of TWA flight 800 are flammable

at these temperatures and at those as low as 96.4° F. Therefore, the Safety Board

concludes that the fuel/air vapor in the ullage of TWA flight 800ís CWT was flammable at

the time of the accident. 

Early in the investigation, questions were raised about whether a Jet A fuel/air

explosion in the CWT, without any additional forces, could generate sufficient pressure to

break apart the fuel tank and lead to the destruction of the airplane. This issue was

explored in July and August 1997 when, as part of a series of tests using an

out-of-service 747 at Bruntingthorpe Airfield, Leicestershire, England, investigators

simulated a fuel/air explosion in the CWT by igniting a propane/air mixture. Evaluation of

the resulting damage indicated that the tank structure failed as a result of overpressure.

The Safety Board recognizes that the test conditions at Bruntingthorpe were not fully

comparable to the conditions that existed on TWA flight 800 at the time of the accident;544

however, previous fuel/air explosions in the CWTs of commercial airliners that had used

Jet A fuel (such as the November 1989 accident involving a 727 operated by Avianca and

the May 1990 accident involving a 737 operated by Philippine Airlines) confirm that a

CWT explosion involving Jet A fuel can break apart the fuel tank and lead to the

destruction of an airplane.

544 For example, those tests were conducted near sea level (rather than at 13,800 feet altitude), did not

account for the effects of aerodynamic loads and pressurization on the accident airplane, used a small charge

as an ignition source, and used a propane/air mixture as a surrogate for Jet A fuel, which released more

energy than the mixture in TWA flight 800ís CWT. 
Analysis 262 Aircraft Accident Report

Further, in the laboratory testing at CIT simulating the altitude, temperature, and

fuel mass loading conditions of TWA flight 800, peak pressures of 39.2 to 52.2 pounds per

square inch (psi) (above ambient pressure)545 were measured from the combustion of a

fuel/air mixture similar to the one that existed in the CWT of TWA flight 800. These

pressures exceed the structural limitations of SWB3 (which, according to Boeing, will fail

when subjected to a pressure differential of between 20 to 25 psi) and, therefore, would be

sufficient to initiate failure of the CWT structure. Finally, analysis of the results of

computer modeling of combustion in a full-scale CWT under conditions simulating those

of TWA flight 800 indicated that a localized ignition of the flammable vapor could have

generated pressure levels that, based upon structural failure analysis, would cause the

damages observed in the wreckage of the accident airplaneís CWT. Accordingly, the

Safety Board concludes that a fuel/air explosion in the CWT of TWA flight 800 would

have been capable of generating sufficient internal pressure to break apart the tank. 

2.2.2 Reported Witness Observations 

As discussed in section 2.2.1.2, the physical evidence indicated that a missile did

not strike the airplane.546 However, numerous witnesses in the vicinity of the accident

reported seeing a streak of light547 before the appearance of a fireball.548 There was intense

public interest in these witness reports and much speculation that the reported streak of

light was a missile that eventually struck TWA flight 800, causing the airplane to

explode.549 However, because the physical evidence indicates that a missile did not strike

the airplane, the witnesses must have been observing something other than a missile. In an

attempt to determine what these witnesses were observing, Safety Board investigators

studied all of the witness accounts in the context of the other information gathered during

this investigation. 

2.2.2.1 Potential Explanations for Reported Witness Observations 

The Witness Group concluded that the streak of light reported by witnesses might

have been the accident airplane during some stage of its flight before the fireball

developed, noting that most of the 258 streak-of-light accounts were generally consistent

with the calculated flightpath of the accident airplane after the CWT explosion. The group

545 At 13,800 feet, the ambient pressure would have been 8.6 psi. The dry bay area in front of SWB3 is

unpressurized and, therefore, at ambient pressure.

546 As will be discussed in section 2.3.1.3, a scenario in which the vapor in the CWT was ignited by a

missile warhead fragment is also very unlikely.

547 According to the witness documents provided to the Safety Board by the FBI, 258 of the

736 witnesses described seeing (about the time of the accident) a streak of light, a flarelike object, or

fireworks in the sky. These and similarly described observations were referred to by the Witness Group as a

ìstreak of light.î For more information about witness observations, see section 1.18.4.

548 Of the 736 witnesses, 599 reported seeing a descending fireball; 200 of those reported seeing the

fireball split into two fireballs before hitting the surface of the water.

549 The witnesses who reported observing a streak of light were the only witnesses who described seeing

something that could have been consistent with a missile. None of the documents pertaining to the other

witnesses provided any information suggestive of a missile.
Analysis 263 Aircraft Accident Report

also determined that the descending fireball observed by 599 witnesses was the accident

airplane and that the 200 witnesses who reported seeing the fireball split into two fireballs

were observing the final stages of the structural breakup of the airplane.

On the basis of computer simulations and witness information, the Safety Board

determined that the entire breakup sequence of the airplane (from the time of the CWT

explosion until the time that the aft portion of the airplane impacted the water) lasted

about 47 to 54 seconds. The sequencing study established that the nose portion of the

airplane separated from the remainder of the airplane after the initial explosion in the

CWT. Computer simulations indicated that this occurred about 3 to 5 seconds after the

initial explosion.550 Computer simulations based on radar data, trajectory calculations, and

airplane performance factors indicate that after the separation of the nose portion, the

remainder of the airplane (including much of the WCS, the wings, the aft fuselage, and the

tail) continued in crippled flight and pitched up while rolling to the left (north), ascended

from 13,800 to about 15,000 or 16,000 feet,551 and then rolled into a descending turn to the

right (south). It is likely that, after the nose portion separated from the aft fuselage, a

fuel-fed fire within the breached CWT (or any other fire that might have existed, such as

from fuel that might have been leaking at the wing roots) would have been visible to

witnesses from some distance and was likely the streak of light reported by many of the

witnesses. 

It is also likely, based on the wreckage, computer simulations, and witness

documents, that shortly after the descending turn to the right, the outboard left and right

wings simultaneously separated at the outboard engines in upward bending. (The recovery

positions of the antennas from the wing tips confirm that the outboard portions of the

wings separated from the airplane significantly after the nose portion.) The separation of

the outboard portions of the wings probably precipitated fuel-fed fires at both outboard

main wing tanks and most likely was the beginning of the developing fireball described by

witnesses. This fireball probably began to develop about 34 seconds after the CWT

explosion.552

On the basis of witness documents, the sequencing study, and computer

simulations, shortly after the outboard portion of the wings separated, the WCS separated

adjacent to the left wing, causing the left wing to separate from what remained of the

airplane structure. These failures would have resulted in the continuing development of a

550 The Safety Boardís computer simulations using airplane performance data indicated that the nose

could have separated from 3.23 to 5.25 seconds after the initial event. (See Main Wreckage Flight Path

Study.) The Safety Board also conducted a trajectory study to calculate how long after the initial event the

nose separated from the airplane. However, because of the uncertainties inherent in such a study (for

example, the effect of winds and ocean currents), this time could not be precisely determined. Results of the

study initially indicated that the nose separation could have occurred from 3.9 to 7.5 seconds after the initial

event. Additional computer simulations conducted after the trajectory study indicated that poor matches

with known radar and wreckage location data were obtained with scenarios in which the nose separated

more than 5 seconds after the initial event. Therefore, the nose separation most likely occurred from about

3 to 5 seconds after the initial event. 

551 Computer simulations indicated that the maximum altitude that the accident airplaneís aft portion

could have ascended after the separation of the nose portion ranged from 15,537 to 16,678 feet. See

addendums 1 and 2 to the Main Wreckage Flight Path Study.
Analysis 264 Aircraft Accident Report

larger fuel-fed fireball.553 The development of a severe fire associated with what remained

of the airplane (portions of the WCS, the right wing, and a few pieces of attached

fuselage) after the separation of the burning left wing is probably what witnesses reported

as a ìsplittingî of the fireball. 

Several witnesses have become prominent because of their unique vantage points

and/or the level of precision and detail provided in their accounts.554 Advocates of the

missile-attack scenario have asserted that some of these witnesses observed a missile.

However, as further explained below, analysis of these witnessesí reported observations

demonstrated that those observations are not consistent with their having observed a

missile attack on TWA flight 800. Rather, they are consistent with these witnesses having

observed some part of the in-flight fire and breakup sequence after the CWT explosion. 

For a missile to have initiated the TWA flight 800 breakup sequence, it would have

had to have been fired at least 41 to 49 seconds before the initial development of a

fireball.555 However, the pilot of the HH-60 helicopter indicated that he saw a fireball

developing just 1 to 5 seconds after he first saw the streak of light.556 Therefore, he must

have been observing the late stages of the airplaneís breakup and not a missile attack.557

The captain of Eastwind Airlines flight 507 also could not have seen a missile attack

because the explosion that he reported having seen occurred about 34 seconds after the

CWT explosion. Therefore, he must have seen the development of the fireball, not the

earlier CWT explosion. Further, the Eastwind captain stated that he did not see a missile or

anything coming up at or hitting the airplane. The CWT explosion also occurred before

the time that the passenger aboard USAir flight 217558 reported seeing a ìflare.î559 In fact,

TWA flight 800 was not in this witnessís field of view at the time of the CWT explosion.

Therefore, this flare could not have been a missile attacking TWA flight 800; however,

this witnessís description is consistent with his having observed burning fuel from the

accident airplane (in crippled flight) and the airplaneís subsequent breakup after the CWT

explosion. 

552 The captain of Eastwind flight 507 reported to the Boston Air Route Traffic Control Center that he

saw an explosion (at 2031:50, about 38 seconds after TWA flight 800 lost power [2031:12] as a result of the

explosion in the CWT). Assuming it took about 4 seconds for the captain to realize that he saw an explosion,

confer with the first officer, and contact ATC, it follows that he first saw the developing fireball about

34 seconds after the explosion. Four seconds is a compromise between what the pilot first reported on the

night of the accident (that he called ATC ìimmediatelyî) and what he stated to Safety Board investigators

almost 3 years later (that he waited about 10 seconds before calling ATC) that favors his initial impression

but provides adequate response time.

553 Many witness accounts describe a fireball developing after the streak of light arched over and began

to descend. Analysis of the witness accounts suggests that the occurrence of this fireball was not a discrete

event, but a transition from a streak of light to a fireball that occurred over a relatively brief period.

554 For more information about these witness accounts, see section 1.18.4.4. 

555 This would comprise a 7- to 15-second missile flight plus about 34 seconds from the time of the CWT

explosion to the outboard wing separations.

556 Although the Safety Board recognizes that witnesses frequently provide inaccurate estimates of time,

the Board considers it highly unlikely that the HH-60 pilot would have reported seeing the streak of light for

only 1 to 5 seconds if he had actually seen the streak of light for 41 to 49 seconds.

557 In addition, the HH-60 pilot stated that the fireball followed the same trajectory as the streak, which is

consistent with the streak and the fireball both representing the airplane.
Analysis 265 Aircraft Accident Report

The witnesses on the boat near Great Gun Beach were close enough to the accident

airplane, with an unobstructed view, to have seen a missile if one had been launched.

However, they did not describe a sequence of events consistent with a missile attack.

Rather, their observations of the northward turn and wing separation are consistent with

their having seen the accident airplane after the CWT explosion. The reported

observations of the witness on the Beach Lane Bridge are also fully consistent with his

having seen the accident airplane from about the time of the CWT explosion until the aft

portion of the airplane descended into the green zone. Further, he did not describe a

sequence of events consistent with a missile strike. Most notably, this witness did not

describe any abrupt change in the streak of lightís speed or direction, which likely would

have been evident if a missile strike had occurred. 

2.2.2.1.1 Anomalous Witness Accounts

Although most of the observations reported in the witness documents are

consistent with the witnesses having observed some portion of the airplaneís crippled

flight and breakup sequence,560 the Safety Board recognizes that a small percentage of the

reported witness observations cannot be explained so readily. The Witness Group noted

that documents pertaining to 38 witnesses described a streak of light ascending vertically,

or nearly so, and that these accounts seem to be inconsistent with the accident airplaneís

calculated flightpath. The Board recognizes that, in addition to these reports, 18 witnesses

reported seeing a streak of light that originated at the surface or from the horizon, which

also does not appear to be consistent with the airplaneís calculated flightpath and other

known aspects of the accident sequence. In an attempt to reconcile these 56 witness

accounts with the calculated accident sequence, the Board considered the effect of

potential deficiencies in the interviewing and documentation process and of errors in

witness memory and/or perception.

558 As noted in section 1.18.4.4, there were six additional witnesses aboard USAir flight 217, one of

whom (seated in seat 6F) stated that he saw an aircraft that ìhad what he took to be its landing lights on, and

[he] could observe cabin lights on the vessel.î According to the document, the witness reported watching the

airplane with ìcabin lightsî for 30 to 40 seconds, then seeing it bank to the left. The document then states,

ìthe aircraft exploded and a large round orange fireball appeared.î The document contains no mention of

any other light in the sky. Because the CWT explosion aboard TWA flight 800 (and the subsequent loss of

electrical power) would have occurred before TWA flight 800 could have been visible to the witness seated

in seat 6F, the passenger in seat 6F could not have seen TWA flight 800ís cabin lights, nor could he have

witnessed any missile attack. Further, his observations are consistent with his having observed the accident

airplane in crippled flight and its subsequent breakup after the CWT explosion.

559 The witness reported seeing the flare about 15 to 20 seconds after he saw a small airplane, which had

to have been the U.S. Navy P-3 Orion, pass under his airplane. The P-3 passed under USAir flight 217 about

the time of the CWT explosion; therefore, if a hypothetical missile or warhead fragment had struck TWA

flight 800, it would have had to have been fired about 7 to 15 seconds before the witness reported seeing the

flare. 

560 This is consistent with the results of the FBI-sponsored line-of-sight study in which a Deputy

Inspector from the Suffolk County Marine Bureau drew a line of sight from the location of 11 witnesses

through a fixed reference point where they each reported having seen a streak of light. (See figure 42.) The

sight lines for 7 of the 11 witnesses studied intersected the green zone (and two passed just north of it),

where the main portion of the wreckage was recovered from the ocean. These sight lines are consistent with

these witnesses observing only the last portion of crippled flight.
Analysis 266 Aircraft Accident Report

2.2.2.1.1.1 Issues Associated with Interviewing and Documentation 

It should be noted that almost all of the witness interviews were conducted by the

FBI during its criminal investigation primarily to determine whether a missile had been

used against TWA flight 800.561 The Safety Board is aware that FBI agents generally did

not record all of the information provided by witnesses but, rather, focused on information

that might have been relevant to the possibility that a missile was used against TWA flight

800.562 This focus is evident in some of the suggested interview questions provided by the

FBI to its agents, which include assumptions about a missile attack.563 Research has

shown that how interview questions are phrased can have a considerable effect on the

information elicited during an interview.564 Research has also shown that other types of

word choices can affect the answers to questions.565

Further, witnesses who believe that their interviewer is knowledgeable about the

event in question are more likely to be influenced by such suggestive questioning than

those who do not think that their interviewer is knowledgeable. Some witnesses may have

believed that the FBI agents were more knowledgeable than they actually were about

whether a missile was used against TWA flight 800. This may have been especially true

during those interviews in which Department of Defense (DoD) missile experts were

present. Therefore, the scope and phrasing of the FBI agentsí questions and possible

witnessesí beliefs that the agents were more knowledgeable about the event than they

actually were (each of which could have suggested the possibility of a missile) could have

biased the interviewees answers to some questions.566

561 As noted in section 1.18.4.1, because of the amount of time that had elapsed by the time the Safety

Board received information about the identity of the witnesses, the Witness Group chose not to reinterview

most of the witnesses; instead, the group chose to rely on the original FBI documents as the best available

evidence of the observations initially reported by the witnesses.

562 During a September 30, 1998, meeting with the Witness Group, an FBI agent who had been involved

with the FBIís earlier criminal investigation indicated that, because of suspicions that a missile might have

been used against TWA flight 800, FBI agents frequently did not record a complete account of the visual and

aural events described by the witnesses but, rather, concentrated on information that appeared relevant to its

criminal investigation. 

563 For example, some of the suggested interview questions included the following: ìHow long did the

missile fly?î; ìWhat does the terrain around the launch site look like?î; and ìWhere was the sun in relation

to the aircraft and the missile launch point?î. For more information about other possible sources of

interviewer or interviewee bias, see section 1.18.4.2.

564 For example, the use of a definite article (for example, ìdid you see the broken headlight?î) instead of

an indefinite article (for example, ìdid you see a broken headlight?î) is sufficient to increase the likelihood

of a positive answer. For more information about this research and other issues associated with the accuracy

of witness memory, see section 1.18.4.3.

565 For example, substantially differing answers were received to the question ìhow fast were the cars

going when they hit each other?î when the word ìhitî was replaced with ìcontacted,î ìbumped,î or

ìsmashed.î

566 For example, one witness described the streak of light as ìgoing straight up, not zig zag.î It is possible

that this witness and others may have used the term ìstraight upî to denote the absence of obvious course

corrections in response to interviewer questions about such corrections, not to indicate that the streak of light

was ascending vertically.
Analysis 267 Aircraft Accident Report

It is also possible that some of the witnesses who were classified ìstreak-of-light

witnessesî by the Witness Group did not, in fact, report observing a streak of light. For

example, a witness may only have observed a fireball falling from the sky; however,

because of deficiencies inherent in the documentation process, the description recorded in

the witness document may have met the Witness Groupís definition of a streak of light.

(The potential for inaccurate witness classification is illustrated by the fact that at least

three streak-of-light witnesses explained during reinterviews by the FBI or media that,

contrary to what was reported in connection with their earlier interview, they did not

observe the streak of light originate at the surface of the earth.) 567

In addition, the Witness Group identified several other issues that made it difficult

to extract accurate and reliable information from the witness documents, including

potentially flawed distance and direction information, multiple witness reports

summarized in a single account, recording of witness speculations and conclusions,

imprecise or vague language, conflicting information, and errors concerning the origin of

the streak of light.568

2.2.2.1.1.2 Issues Associated with Witness Perception and Memory

Perception569

It is possible that, from some witnessesí locations, the burning airplane might have

appeared to be rising vertically. Specifically, some witnesses who were standing on Long

Island looking to the southwest would have been in a position to see the airplane coming

nearly directly toward them, which could have resulted in the appearance of a rising

(perhaps even near-vertical to some witnesses) flightpath for a portion of the time that the

airplane was in crippled flight. Because of the angle at which these witnesses were

positioned, any fire coming from the CWT could have appeared to them as a streak of

light moving vertically.

It should also be noted that witnesses observing the accident from Long Island

would have been viewing it against a featureless background and, therefore, would have

had few, if any, depth perception cues. Consequently, because of the tendency to perceive

more distant objects as being about the same distance away as intervening objects (the

equidistance tendency), the streak of light could have appeared closer to them than it

actually was. The equidistance tendency is illustrated in figure 45, which depicts a person

at point ìcî observing an object at point ìaî that appears at point ìbî at the same

perceptual distance as an intervening building at point ìd.î 

This phenomenon possibly explains why the witness on the Beach Lane Bridge

stated that the streak of light came into his view just above the roof of a house and why he

reported that it appeared to have originated on the beach just beyond that house. For some

observers, the equidistance tendency would have caused the streak of light to appear lower

567 For further information about these three witnesses, see section 1.18.4.2.

568 For more information about these issues, see section 1.18.4.2.

569 For more information about perception of distance and motion, see section 1.18.4.3.1.
Analysis 268 Aircraft Accident Report

in the sky than it actually was, which explains why the streak of light may have been

reported by some witnesses as having originated at the surface of the earth. It should be

noted that almost all of the witnesses who reported that the streak of light originated at the

surface viewed the streak of light on the far side of intervening terrain features, such as

houses or the barrier island. These witnesses would have perceived the streak of light to be

about the same distance as the intervening object because the elevation angle of the object

would have been perceptually reduced.

Further, if an observer views an object for which few or no depth perception cues

are available, the motion of the observerís head will cause the observer to perceive the

object moving concurrently with the observerís head. At the time of the reported

observations, 13 of the 38 witnesses who reported that the streak of light ascended

vertically or nearly so were in vehicles (including boats) and one of the witnesses was

paddling a surfboard. Other witnesses may also have been moving. The effect of the

concurrent motion of the observerís head and the observed object would have caused the

streak of light to appear to have been moving differently than it was actually moving, and

this effect could have caused the streak of light to have appeared to rise vertically or nearly

so. The closer that the streak of light appeared to a given witness, the greater the apparent

motion of the streak of light to that witness.

Further, the airplaneís relatively low elevation above the horizon570 could have led

many witnesses to believe that they were observing a flare or fireworks. (As discussed

previously, because of the absence of depth perception cues, witnesses who saw the

airplane in crippled flight could not be expected to provide good information about how

far away it was or its direction of travel.)571 As discussed in section 1.18.4.3, a witnessís

perception of events is driven by context cues, experience, and expectations. The fact that

it was July (a month often associated with fireworks) and that several flares had been

launched in the area during the week preceding the accident could have led witnesses to

570 For example, to witnesses located 10 nm, 15 nm, or 20 nm (slant range) from the accident airplane,

the airplane would appear about 13.1o, 8.7o, and 6.5o, respectively, above the horizon.

Figure 45. A graphical illustration of the equidistance tendency.
Analysis 269 Aircraft Accident Report

believe that they were seeing a flare or fireworks, which could also have led witnesses to

believe, and subsequently report, that what they observed had originated from the surface,

even though they did not actually see the objectís origination point. It is also possible that

some of these witnesses may have actually observed one or more of the flares launched

from boats in the area shortly after the wreckage impacted the water to direct attention to

the accident site.572

Memory573

Research has shown that a person may unintentionally, and without awareness,

create an inaccurate memory because of exposure to postevent information provided by

other witnesses or media reports or because of expectations and beliefs. If misinformation

is received from one of these sources, there is an increased likelihood that this

misinformation will be assimilated and reported by the witness as a memory. Some

witnessesí recollections might have been influenced by exposure to other witnessesí

accounts; being interviewed together; engaging in conversations with other people; or

hearing, watching, or reading media accounts of the crash (which included interviews with

eyewitnesses and speculation about a missile strike.)574 In this regard, it is noteworthy that

the documents recording the initial interviews of some witnesses did not indicate that

these witnesses initially reported seeing a missile; however, these same witnesses have

since indicated that they believe that a missile was used against TWA flight 800. Research

has also shown that a witnessís confidence in the accuracy of his or her memory (even

when it is inaccurate) increases over time and with repeated retrieval, which possibly

explains why some witnesses displayed a high level of confidence in the accuracy of their

reported observations. 

2.2.2.2 Conclusion Regarding Reported Witness Observations

Given the potential deficiencies in the witness documents, the potential

inaccuracies in witnessesí recollections, and the large number of witnesses whose

observations were documented and studied in this case, the Safety Board did not expect all

of the documented witness observations to be consistent with one another. Therefore, the

571 This is illustrated by the variation in the witness accounts regarding the direction of the motion of the

streak of light (38 reported that it moved vertically; 45 reported that it moved to the east; 23 reported that it

moved to the west; 18 reported that it moved to the south; and 4 reported that it moved to the north.

Thirty-six of the streak-of-light accounts contained no information about the direction of the motion of the

streak of light). 

572 Although flares were launched after the accident sequence (and, therefore, would have appeared after

the fireball), witnesses who also reported seeing a fireball may have recalled seeing the streak of light (flare)

before the fireball because their memories of the accident sequence were influenced by postevent

information provided by media reports or other witnesses or by other factors. (As discussed in

section 1.18.4.3.2, research has shown that when asked to recall information about an event, people

sometimes recall information about an event that occurred at a different time than the event about which

they are questioned.) 

573 For more information about storing information in and retrieving it from memory, see

sections 1.18.4.3.1 and 1.18.4.3.2.

574 Many witnesses acknowledged that they did not even know that they had observed an airplane

accident until after seeing media accounts of the accident. 
Analysis 270 Aircraft Accident Report

Board did not view these apparently anomalous witness reports as persuasive evidence

that some witnesses might have observed a missile. 

Investigators determined that if witnesses had observed an actual missile attack on

TWA flight 800 (beginning about the time that an airborne missile would have become

visible to the time that the wreckage from TWA flight 800 fell into the ocean), they would

have seen the following: (1) a light (the burning of the missile motor) ascending very

rapidly and steeply for about 8 seconds (this rapidly moving light, which would have been

visible for at least 12 nm575 from the launch point, would not descend like a firework or

flare); (2) the light disappearing576 for up to about 7 seconds;577 (3) upon a missile (or

warhead fragment) striking the airplane and igniting the fuel/air vapor in the CWT,

another light (flames coming from the CWT), moving considerably slower and more

laterally than the first light (although ascending somewhat for a short time), for about

30 seconds; (4) this light descending while simultaneously developing into a fireball

falling toward the ocean. It is noteworthy that none of the witness documents included a

description of such a scenario.

Therefore, the Safety Board concludes that the witness observations of a streak of

light were not related to a missile and that the streak of light reported by most of these

witnesses was burning fuel from the accident airplane in crippled flight during some

portion of the postexplosion, preimpact breakup sequence.578 The Board further concludes

that the witnessesí observations of one or more fireballs were of the airplaneís burning

wreckage falling toward the ocean.

575 Although during some of the missile visibility tests the light from the burning missile motor was

visible to witnesses who were positioned up to 14 nm from the launch point, the lighting conditions during

those tests were darker than in earlier tests and, therefore, were not comparable to the lighting conditions at

the time of the TWA flight 800 accident. For the tests conducted under lighting conditions that were

comparable to those that existed at the time of the TWA flight 800 accident, no witnesses were positioned

farther than 12 nm away from the launch site.

576 The Safety Boardís missile visibility tests indicated that, after completion of the motor burn,

observers saw no visible evidence of the missile until it selfdestructed.

577 Because of the distance from which a hypothetical missile would likely have been launched, most

direct-strike scenarios and any self-destruct scenario (for further discussion of a missile self-destruct

scenario, see section 2.3.1.3) would require that the missile travel about 7 seconds in ìcoastî mode;

therefore the missile would have been invisible before striking the airplane. Only if a missile had been

launched from almost directly underneath the accident airplane (as close to vertically as possible) would the

ìcoastî time have been significantly reduced or eliminated. 

578 Central Intelligence Agency (CIA) analysts, who initially evaluated witness information provided by

the FBI, also concluded that the majority of the witness observations must have occurred well after the

initial explosion and, consequently, that none of those observations could have been of a missile causing the

explosion. The CIA analysts concluded that the remaining witness accounts described events fully

consistent with observations expected if only the airplane were being observed in various stages of crippled

flight. The CIA analysts concluded that there was absolutely no evidence that a missile was employed to

shoot down TWA flight 800.
Analysis 271 Aircraft Accident Report

2.2.3 Cause of the In-Flight Breakup

On the basis of the accident airplaneís breakup sequence; wreckage damage

characteristics; scientific tests and research on fuels, fuel tank explosions, and the

conditions in the CWT at the time of the accident; and analysis of witness information, the

Safety Board concludes that the TWA flight 800 in-flight breakup was initiated by a

fuel/air explosion in the CWT. 

2.3 Evaluation of Potential Ignition Sources 

In an attempt to determine what ignited the flammable fuel/air vapor in the CWT

and caused the explosion, investigators evaluated numerous potential ignition sources.

Possible ignition source scenarios fell generally into two main categories: (1) energy

entered the CWT by way of fuel quantity indication system (FQIS) wiring (discussed in

section 2.3.2) or (2) energy entered the CWT by other means (discussed in section 2.3.1).

The minimum ignition energy (MIE) for Jet A fuel has been generally accepted to

be about 0.25 millijoule (mJ)579 based on testing done by the American Petroleum Institute

(API) and is documented in API Recommended Practice 2003 (API 2003), ìProtection

Against Ignitions Arising out of Static, Lightning, and Stray Currents,î fifth edition,

December 1991. However, laboratory testing conducted at CIT using fuel similar to that

used on the accident flight indicated that the energy needed to ignite Jet A fuel at an

altitude of 13,800 feet ranged from 0.5 mJ (at 122° F) to 500 mJ (at 104° F). Therefore,

this higher range of ignition energies was used as a reference for evaluating whether a

particular potential ignition source could have ignited the fuel/air vapor in TWA flight

800ís CWT. However, investigators also considered the lower, generally accepted, MIE of

0.25 mJ to be relevant in evaluating whether a potential ignition source might pose an

ignition hazard in other airplanes under different circumstances (for example, at different

altitudes and temperatures). 

2.3.1 Possible Ignition Sources Unrelated to the Fuel Quantity 

Indication System 

The Safety Board evaluated numerous potential ignition sources that could have

ignited the fuel/air vapor in TWA flight 800ís CWT. Among the potential ignition sources

considered by investigators that were unrelated to the FQIS were the following: a

lightning or meteorite strike; a missile fragment; a small explosive charge placed on the

CWT; auto ignition or hot surface ignition, resulting from elevated temperatures produced

by sources external to the CWT; a fire migrating to the CWT from another fuel tank via

the vent (stringer) system; an uncontained engine failure or a turbine burst in the air

conditioning packs beneath the CWT; a malfunctioning CWT jettison/override pump; a

malfunctioning CWT scavenge pump; and static electricity. 

579 A mJ is one-thousandth of a joule (J), which is a measurement of electrical energy (1 J is the work

done by 1 watt in 1 second).
Analysis 272 Aircraft Accident Report

2.3.1.1 Lightning Strike

No witnesses reported that lightning struck the accident airplane or that lightning

was in the area of TWA flight 800 at the time of the accident. Further, recorded

meteorological data indicated no atmospheric disturbances in the area of TWA flight 800

at the time of the accident. Therefore, it is very unlikely that a lightning strike ignited the

fuel/air vapor in the CWT.

2.3.1.2 Meteorite Strike

An expert in the study of meteorites testified at the Safety Boardís public hearing

that no report exists of a meteorite ever having struck an airplane. He testified that, given

the known frequency with which meteorites have struck cars and dwellings over the past

several decades and based on a comparison of estimates of (1) the area of the earthís

surface represented by cars and dwellings with (2) the area of the earthís surface

represented by airplanes in the air at any given time, he calculated that a meteorite could

be expected to strike an aircraft once every 59,000 to 77,000 years. Further, none of the

holes in the CWT or fuselage structure exhibited characteristics of high-velocity

penetration through the exterior (such as outward metal deformation around the perimeter

of the entry and exit sides of the hole, lack of overall deformation, and distinctive features

of the hole wall surface) that would be expected from a meteorite strike.580 Therefore, it is

very unlikely that a meteorite ignited the fuel/air vapor in the CWT. 

2.3.1.3 Missile Fragment

Although, as discussed in section 2.2.1.2, a missile-strike scenario was not

supported by the physical evidence, investigators considered whether a missile might have

selfdestructed close enough to TWA flight 800 for a missile fragment to have entered the

CWT and ignited the fuel/air vapor, yet far enough away that it would not have left any

damage features characteristic of a missile strike. According to missile performance data

provided to the Safety Board by the Naval Air Warfare CenteróWeapons Division

(NAWC-WD), Survivability/Vulnerability Information Analysis Center, and Missile and

Space Intelligence Center (MSIC), the warhead from a shoulder-launched missile will

selfdestruct about 15 seconds after launch if no target has been hit,581 and missile warhead

fragments are always dispersed in a predictable fragmentation pattern. Using the initial

speed and decay rate for the missile warhead fragments, the Board calculated that a

detonation would have to occur within 40 feet of an airplane for a fragment traveling

perpendicular to the CWT to retain sufficient velocity to penetrate the CWT. In the case of

a fragment traveling at an angle to the CWT, a detonation would have to occur even closer

to the airplane.

580 Two holes in the horizontal pressure deck (above the left side of the wing landing gear bay) contained

some features of both high-velocity and low-velocity penetrations. However, the physical features of both

holes indicated that they were made by objects traveling from inside the airplane out, rather than the reverse.

Therefore, neither hole could have been made by a meteorite.

581 The exact number of seconds after launch that a missile warhead will selfdestruct varies slightly,

depending on the type of missile. The Safety Boardís analysis of the likelihood of this scenario would apply

equally to missiles with a slightly longer or shorter self-destruct range.
Analysis 273 Aircraft Accident Report

The Safety Boardís computer simulations using this missile performance data,

which took into account the winds at the time of the accident, showed that the number of

locations on the surface (all of which were on water) from which a missile launched at

TWA flight 800 could have resulted in selfdestruction within 40 feet of the airplane were

extremely limited. Simulations further showed that, because of the pattern in which

missile warhead fragments are dispersed after selfdestruction, launches from many, if not

most, of these positions would not result in fragments with a proximity and orientation

sufficient to penetrate the CWT. Further, simulations also showed that a missile launch

from one of the few sites that could potentially have resulted in a detonation at the

proximity and orientation for a fragment to penetrate the tank and ignite the fuel/air vapor

would have resulted in numerous fragments impacting the surrounding airplane structure

with sufficient force to leave distinctive impact marks.582 No such impact marks were

found on the accident airplaneís wreckage.

In addition, a missile fragment would have had to penetrate multiple layers of

material to enter the CWT, and, although some areas of fuselage skin and CWT structure

are missing, possible straight-line entry paths were extremely limited. Further, as

previously mentioned, none of the holes in the CWT or fuselage structure exhibited

characteristics of high-velocity penetration through the exterior that would have resulted

from a missile fragment.583 Therefore, it is very unlikely that a missile warhead

selfdestructed at a distance close enough to the accident airplane that at least one fragment

could have entered the CWT with sufficient force and velocity to have ignited the fuel/air

vapor in the CWT, yet far enough from the airplane that almost none of the remaining

fragments impacted its structure.584

2.3.1.4 Small Explosive Charge

Another potential ignition source that investigators considered was the detonation

of a small explosive charge585 that had been placed on the CWT that could have breached

the fuel tank and ignited the flammable fuel/air vapor. However, testing by the Safety

Board and the British Defense Evaluation and Research Agency in August 1997 at

Bruntingthorpe Airfield, Leicestershire England, demonstrated that when metal of the

582 According to data from MSIC, upon detonation a warhead will release about 1,000 fragments of

varying sizes, up to 40 grains (equivalent to 0.0916 ounce).

583 The features of the two holes in the horizontal pressure deck that displayed some features of

high-velocity penetration indicated that they were made by objects coming from different directions;

therefore, it is not likely that they originated from a common location. Further, as previously mentioned, the

physical features of both holes indicated that they were made by objects traveling from inside the airplane

out, rather than the reverse. Therefore, neither hole supports a scenario in which the fuel/air vapor in the

CWT was ignited by the entry of a missile fragment.

584 A study by the NAWC-WD for the DoDís Office of Special Technology also concluded that the

probability of such a scenario was ìextremely low.î According to the report, ì[f]or a missile shooter to

launch the weapon, the weapon to fly at the target, the target to be just beyond the kinematic envelope of the

weapon, the warhead to be properly positioned at self-destruct, and one of the very few heavy fragments to

impact in an area to penetrate the thick wing skin with enough residual energy to initiate an explosion is

stretching credulity to the point that this scenario must be considered a low-priority area of concern.î

585 A small explosive charge is a minimal amount of a highly explosive material (such as a plastic

explosive) detonated by a fusing device.
Analysis 274 Aircraft Accident Report

same type and thickness as the CWT walls was penetrated by a small charge, there was

petalling of the surface on which the charge was placed, pitting on the adjacent surfaces,

and visible hot gas washing damage in the surrounding area. As discussed in

section 2.2.1.2, none of these damage characteristics were found on the recovered CWT

wreckage, and none of the areas of missing skin were large enough to represent the

damage that would have been caused by the detonation of a small explosive charge.

Therefore, it is very unlikely that a small explosive charge detonated on the CWT,

breached the fuel tank, and ignited the fuel/air vapor. 

2.3.1.5 Auto Ignition or Hot Surface Ignition

Investigators also evaluated whether the surface of the CWT might have been

heated to a high enough temperature to have caused ignition of the fuel/air vapor inside

the tank as a result of auto ignition586 or hot surface ignition.587 Investigators considered

whether an external mechanism, such as a bleed air leak near the CWT, a fire in the air

conditioning pack bay beneath the CWT, or a fire in the main landing gear wheel well,

could have raised a large volume of the CWT ullage above the auto ignition temperature

or caused a localized hot spot on the interior of the CWT skin that could have reached hot

surface ignition temperatures.

However, a scenario in which a significant portion of the CWT volume was heated

to auto ignition temperatures was not considered a credible possibility because the amount

of heat that would have been necessary to raise the temperature of the entire (or at least

most of the) CWT to 460° F or more would have almost certainly left significant evidence

of thermal damage on the CWT.588 Further, for a hot spot approaching hot surface ignition

temperatures (900° to 1,300° F) to have developed on the inside surface of the CWT, the

temperature on the outside surface would have had to have been at least as high as these

temperatures. However, these temperatures are near the softening and melting

temperatures of the CWTís aluminum skin; therefore, the CWT would have shown

evidence of having been heated to these temperatures. As further discussed below, no

evidence was found that indicated that the CWT reached auto ignition or hot surface

ignition temperatures. 

Analysis of Boeing data indicates that multiple engine malfunctions could produce

bleed air temperatures above 800° F at the engines. However, the accident airplaneís

586 Auto ignition can occur when a large volume of Jet A fuel/air vapor is sufficiently elevated in

temperature without direct contact with an ignition source, such as a spark, arc, or hot surface or filament.

Auto ignition is highly dependent upon many factors, including the size of the heated volume and container

and other environmental conditions that affect fuel vapor. Existing research indicates that the auto ignition

temperature at sea level for Jet A fuel is about 460° F. Auto ignition temperature increases as the altitude

increases. 

587 Hot surface ignition can occur when a very high temperature surface comes into contact with Jet A

fuel or fuel/air vapor. Hot surface ignition is highly dependent upon many factors, including the geometric

aspects of the hot surface and other environmental conditions that affect fuel and fuel/air vapor. Existing

research indicates that hot surface ignition temperatures at sea level for fuels similar to Jet A range from

900° to 1,300° F. Hot surface ignition temperature increases as the altitude increases.

588 Knowledge of the CWT thermodynamic environment, derived in part from the JFK flight tests and

the Marana, Arizona, ground tests, indicates that such temperatures would most likely not be reached.
Analysis 275 Aircraft Accident Report

engines, all of which were recovered and extensively examined, showed no evidence of

any failure or malfunction. Further, engineering data from Hamilton Standard (the

manufacturer of the air conditioning system) and Boeing indicated that the combination of

failures required to produce such elevated bleed air temperatures589 would also have

caused an overpressure leading to an air conditioning duct burst. However, the

reconstructed air conditioning packs showed no evidence of leakage in the packs or ducts.

In addition, no discussion of an overheat or other warning590 was recorded by the CVR,

which would be expected if excessive bleed air temperatures or any other failure occurred,

especially with two flight engineers in the cockpit.591

Investigators also considered whether a fuel leak from the CWT or the inboard fuel

tanks onto the hot air conditioning packs beneath the CWT could have caused a fire in the

pack bay or main landing gear wheel well and subsequent auto ignition or hot surface

ignition of the fuel/air vapor in the CWT. (The Safety Board is aware of at least one report

of fuel leaking into the air conditioning pack bay in a 747.) However, no fire damage was

observed on the first two packs (and surrounding material) to depart the airplane (those

found in the red zone),592 nor on the fairings from underneath the packs, which would be

expected if the packs had been exposed to a fire of sufficient heat and intensity to have

raised a large volume of the temperature inside the CWT to 460° F or to have heated a spot

on the interior surface to more than 900° F. Although there was fire damage to the wheel

well, the burning and soot in that area had a flow pattern that could be traced from the fuel

tank to the wheel well, indicating that the fire in the CWT preceded the fire in the wheel

well. Further, the tires from the wheel well showed only limited localized heat damage,

contrary to what would be expected if they had been exposed to a fire of sufficient heat

and intensity to have raised a large volume of the temperature inside the CWT to 460° F or

to have heated a spot on the interior surface to more than 900° F. Therefore, it is very

unlikely that the fuel/air vapor in the CWT ignited as a result of either auto ignition or hot

surface ignition caused by elevated temperatures produced by sources external to the

CWT. 

2.3.1.6 Fire Migration Through the Vent System

Because the fuel vent system on the 747 permits communication between fuel

tanks, investigators considered the possibility that a fire originating in one of the other fuel

tanks might have entered the CWT through the fuel vent system. However, the fuel vent

systemís design is such that if a flame had propagated through the vent system, it would

589 Concurrent failures of the (1) engine bleed air pressure relief valve, (2) precooler cooling airflow

control valves, and (3) pylon shutoff and regulating valve would be required.

590 The failures required to produce such excessive bleed air temperatures would have activated engine

malfunction and excessive bleed air temperature warnings from sensors installed both inside and outside the

bleed air ducts.

591 It is evident from the captainís remark on the CVR about the ìcrazyî No. 4 fuel flow that the flight

crew was attentive to the gauges and other indications on the instrument panel.

592 The third pack, which was recovered from the green zone, had fire damage. However, this damage

was likely the result of the severe fire(s) that affected the WCS and surrounding structure during the

remainder of the breakup sequence, most notably after the separation of the left wing. 
Analysis 276 Aircraft Accident Report

have had to have traveled through the wing tip surge tank to meet the vents from the other

fuel tanks. The surge tank overboard vents are equipped with a surge tank protection

(STP) system, which includes a photocell in each overboard vent designed to trigger the

discharge of Halon (a fire extinguishing agent) into the surge tank when a flame or bright

light source is sensed at an overboard vent. The accident airplaneís Halon bottles were

found undischarged.593 More importantly, temperature measurements taken during the

1997 flight tests at JFK showed that at 13,800 feet, the fuel/air vapor in the section of the

vent system leading to the wing tips would have cooled to the point that it could not

sustain a flame. Therefore, it is very unlikely that a fire originating in another fuel tank

migrated through the fuel vent system to the CWT and ignited the fuel/air vapor.

2.3.1.7 Uncontained Engine Failure or Turbine Burst

Investigators also considered the possibility that an uncontained engine failure or a

turbine burst in the air conditioning packs beneath the CWT caused debris to enter the

tank with sufficient velocity to ignite the vapors. However, the engines and reconstructed

packs showed no evidence of an uncontained engine failure or turbine burst. Therefore, it

is very unlikely that an uncontained engine failure or a turbine burst in the air conditioning

packs ignited the fuel/air vapor in the CWT.

2.3.1.8 Malfunctioning Jettison/Override Pump

Investigators also considered the possibility that a jettison/override pump

malfunction produced a spark that ignited the fuel/air vapor in the CWT. Both pumps from

the CWT were recovered and tested. One pump had evidence that indicated that the

impeller had contacted the impeller housing. However, tests conducted in connection with

the investigation of the May 11, 1990, accident involving a CWT explosion on a

Philippine Airlines 737 indicated that galling of the aluminum impeller against the

impeller housing would not result in ignition of flammable vapor. The other pump showed

no evidence of any malfunction or overheating that could have caused the vapor in the

pump to ignite.

There were no comments or sounds on the CVR indicating that the pumps were

used (and there would have been no reason for them to have been used because there was

no usable fuel in the CWT). However, even if the pumps were used and a pump motor had

malfunctioned or overheated, this would not likely have ignited the fuel/air vapor in the

CWT because of the flame suppression system incorporated into the pump design. When

vapors in the motor housings of both an old and a new scavenge pump (which have a

flame suppression system similar to that in the jettison/override pumps) were ignited in

postaccident testing, the test pump flame suppression systems consistently functioned as

designed and prevented sparks or flames from escaping from the pump motors and

igniting flammable vapors outside the pumps. (In some tests, even after investigators

593 It should be noted that the STP system is designed, and the Halon bottles are configured, to protect

against flames entering the airplane from the outside. Therefore, even though the passage of a flame through

the surge tank would likely have triggered discharge of the Halon, the flames would not likely have been

extinguished by the Halon.
Analysis 277 Aircraft Accident Report

removed part of the flame suppression system to simulate conditions found in some

malfunctioning pumps,594 the remaining parts of the system still prevented ignition of the

surrounding flammable vapor.) Further, in previous instances in which arcing has occurred

in jettison/override pump motors, the flame suppression system has prevented a flame

from entering the fuel tank and igniting the fuel/air vapor.595 Therefore, it is very unlikely

that a jettison/override pump malfunction ignited the fuel/air vapor in the CWT.

2.3.1.9 Malfunctioning Scavenge Pump 

Investigators also considered the possibility that a scavenge pump malfunction

could have produced the ignition energy for the CWT explosion. The motor and impeller

were not recovered and, therefore, could not be examined for signs of fire or overheating.

However, the recovered control switch had damage indicating that it was in the ìOffî

position at impact, suggesting that the pump was not operating at the time of the explosion

and, therefore, could not have provided an ignition source. However, even if the pump had

been on596 and a flame had developed inside the pumpís motor housing, it is still unlikely

that the flame could have entered the tank.597

Although examination of a section from the recovered scavenge pump inlet tube

(which draws fuel from the CWT into the pump) revealed some darkened areas on the

tube, there was insufficient evidence to determine whether this damage occurred as part of

the initial event or later in the accident sequence. However, as previously discussed, when

vapors in the motor housings of both an old and a new scavenge pump were ignited in

postaccident testing, the test pump flame suppression systems consistently functioned as

designed and prevented sparks or flames from escaping from the pump motors and

igniting flammable vapors outside the pumps.

Examination also found no evidence of extreme localized overheating (such as

heat damage or discoloration to the flange/spacer between the pump and the rear spar or

splatter on surrounding structure from a hole arced through the motor housing), which

could have resulted in ignition of the fuel/air vapor inside the CWT. Therefore, although a

CWT scavenge pump malfunction could not be entirely ruled out as a possible ignition

source for the CWT explosion, there is no evidence that such a malfunction ignited the

fuel/air vapor in the CWT. 

594 The flame suppression system consists of a small-diameter tube in series with a check valve. Tests

were run with one or the other component removed.

595 For example, 18 jettison/override pumps that had experienced arcing in the motor were examined by

Boeing for evidence of sulfide deposits. (See Systems Group Chairmanís Factual Report Addendum,

ìDeposits on Fuel Tank Electrical Components.î) None of these pumps had ignited the vapor in the pump.

596 Investigators recognized the possibility that the switch might have moved to the ìOffî position

sometime during the breakup sequence. However, there were no comments on the CVR indicating that the

flight crew was activating the CWT scavenge pump. Such an action would have been highly unusual given

that the tank was not in use and, therefore, would likely have generated some comment from one or both of

the flight engineers.

597 The scavenge pump motor is mounted on the outside of the CWT, on the aft of the rear spar, and

cooling fuel is drawn from the tank through small-diameter tubes.
Analysis 278 Aircraft Accident Report

2.3.1.10 Static Electricity

Investigators also theorized that electrically isolated parts in the CWT598 could

become so highly charged from static electricity (generated by fuel striking the part or

from fuel moving through the fuel lines) that a discharge between such a part and a nearby

grounded fuel tank component could create a spark of sufficient energy to ignite the

fuel/air vapor in the tank. The Safety Board contracted with the U.S. Air Force (USAF)

Research Laboratory (AFRL) at the Wright Patterson Air Force Base and with the Naval

Research Laboratory to conduct tests to assess the electrostatic charging capabilities of

several electrically isolated components commonly found in aircraft fuel systems. This

testing revealed that unbonded Teflon-cushioned wire clamps (found in the 747 CWT and

inboard main fuel tanks) were particularly susceptible to electrostatic charging. However,

the highest voltage potential attained in testing that used fuel similar to that used on the

accident flight was 650 volts, which it was estimated would produce a discharge energy of

only about 0.0095 mJ.599 Even assuming that the clamp could attain a voltage of

1,150 volts (which was produced in tests using fuel of a higher conductivity than the fuel

in the CWT on TWA flight 800), the highest discharge energy that voltage could produce

was estimated to be about 0.030 mJ, which is still well below the 0.25 mJ MIE for Jet A

fuel vapor. Therefore, on the basis of the evidence from the TWA flight 800 investigation

and research and testing to date, it is very unlikely that static electricity ignited the fuel/air

vapor in the CWT. Nonetheless, the Board recognizes that the discharge of static

electricity has resulted in prior fuel tank explosions600 and cannot rule out the possibility

that, under some conditions, static electricity could produce sufficient energy to ignite the

flammable vapor in a fuel tank. 

The findings from this investigation raised concerns regarding the adequacy of

bonding protection against high-energy discharges, such as from lightning (or perhaps

from refueling). The Safety Board is concerned about the Boeing design practice that

permits parts that are less than 3 inches long in any direction (including some types of

clamps and connectors installed in fuel tanks) to be electrically unbonded. Tests conducted

by Boeing after two fuel tank explosions in 1970 found that a single unbonded clamp

could store up to 5 mJ of energy during the increased fuel flow rates associated with

ground refueling.601 Further, according to the FAAís Aircraft Lightning Protection

Handbook, unbonded clamps could present an ignition hazard in the event of a lightning

598 Certification standards only require that ìmajor componentsî of the powerplant installation be

electrically bonded to the other parts of the airplane. See 14 Code of Federal Regulations

(CFR) 25.901(b)(4). According to a June 2, 2000, letter to the Safety Boeing, Boeing design practices permit

parts that are less than 3 inches long in any direction (including some types of clamps and connectors

installed in fuel tanks) to be electrically unbonded, presumably because such parts are not believed to have

enough capacitance to retain hazardous levels of static electricity under expected operating conditions. 

599 A volt is the basic unit of measurement of electromotive force (the force that causes electrons to flow

through a conductor). As previously mentioned, a J is a measurement of electrical work or energy. Power

(watts or J/seconds) is related to voltage through V2/R, or the square of the voltage divided by the resistance,

measured in ohms. 

600 The July 23, 1998, report of the Aviation Rulemaking Advisory Committee (ARAC) Fuel Tank

Harmonization Working Group (FTHWG) and API 2003 cite several previous fuel tank explosions believed

to have been caused by static electricity. 
Analysis 279 Aircraft Accident Report

strike. The Safety Board concludes that Boeingís design practice that permits parts less

than 3 inches long in any direction to be electrically unbonded may not provide adequate

protection against potential ignition hazards created by static electricity or electricity

generated by lightning. Because the Board is concerned that other manufacturersí design

practices may not be adequate, the Safety Board believes that the FAA should examine

manufacturersí design practices with regard to bonding of components inside fuel tanks

and require changes in those practices, as necessary, to eliminate potential ignition

hazards. 

2.3.1.11 Conclusion Regarding Ignition Sources Unrelated to the Fuel 

Quantity Indication System

In summary, the Safety Board concludes that it is very unlikely that the flammable

fuel/air vapor in the CWT on TWA flight 800 was ignited by a lightning or meteorite

strike; a missile fragment; a small explosive charge placed on the CWT; auto ignition or

hot surface ignition, resulting from elevated temperatures produced by sources external to

the CWT; a fire migrating to the CWT from another fuel tank via the vent (stringer)

system; an uncontained engine failure or a turbine burst in the air conditioning packs

beneath the CWT; a malfunctioning CWT jettison/override pump; a malfunctioning CWT

scavenge pump; or static electricity. 

2.3.2 Introduction of Ignition Energy Through the Fuel Quantity 

Indication System

The only electrical wiring located inside the CWT is the wiring associated with the

FQIS.602 According to Boeing design specifications, the voltage to the FQIS wiring is

limited so that it cannot discharge energy in excess of 0.02 mJ. Therefore, for the FQIS to

have played a role in igniting the flammable fuel/air vapor in the CWT, the following two

events would have had to have occurred: (1) a transfer of a higher-than-intended voltage

onto FQIS wiring from a power source outside of the fuel tank and (2) the release of the

energy from that FQIS wiring into the inside of the tank in a way that could ignite the

fuel/air vapor in the tank.

2.3.2.1 Transfer of Energy to the Fuel Quantity Indication System Wiring 

Outside the Center Wing Tank

2.3.2.1.1 Electromagnetic Interference

Investigators examined several possible ways by which energy (from either

outside or inside the airplane) might have been transferred to the FQIS by electromagnetic

interference (EMI). One possibility explored was that radiated electromagnetic energy

601 These test results were reported to the FAA during a March 11, 1971, meeting regarding electrostatics

during refueling and were summarized in a March 16, 1971, FAA memorandum that documented the results

of that meeting.

602 Although the fuel pump motor cavities contain wiring, they are separated from the CWT by flame

suppression passages and other parts.
Analysis 280 Aircraft Accident Report

from radio frequency (RF) transmitters outside the airplane could couple to FQIS wiring.

However, the National Aeronautics and Space Administrationís (NASA) Langley

Research Center (LaRC) study of the RF sources present in the electromagnetic

environment (EME) at the time of the TWA flight 800 accident concluded that, even

assuming that the highest possible electromagnetic signal levels were being emitted by

these sources at the time of the accident, the maximum available energy inside the airplane

(but not on the wires) from these signal levels would be less than 0.1 mJ. The study further

noted that any such energy that coupled to aircraft wiring inside the airplane would be

several orders of magnitude less than 0.1 mJ. Therefore, although it is conceivable that

under different circumstances a significantly stronger EME than was present at the time of

the TWA flight 800 accident might present an ignition hazard, the Safety Board concludes

that EMI from RF sources external to TWA flight 800 did not produce enough energy to

ignite the fuel/air vapor in the CWT.

The Safety Board also considered the possibility that power was transferred from

personal electronic devices (PED) to the FQIS wiring. Therefore, at the Boardís request,

NASA LaRC scientists also studied the electromagnetic force that could be generated by

PEDs. The study found that the strongest electrical signal that could be coupled to the

FQIS wiring from PEDs was about one-tenth of the minimum power level required in

laboratory tests to induce a spark from the FQIS by applying a simulated arc gap.603

Therefore, the Safety Board concludes that EMI from PEDs played no role in igniting the

fuel/air vapor in TWA flight 800ís CWT.

Finally, the Safety Board examined the possibility that a transient voltage/current

spike from wiring associated with higher-voltage aircraft systems could be induced onto

corouted FQIS wiring. Early in the investigation during laboratory tests, Boeing was able

to induce transient voltages in excess of 1,000 volts (estimated to be capable of producing

0.6 mJ of energy) in FQIS wiring. However, these tests did not accurately represent the

voltage sources that would exist in an actual airplane. More comprehensive testing was

conducted by the Naval Air Warfare CenterñAircraft Division on an airplane using actual

airplane systems, in which transient voltages were created in an energized FQIS by

switching the power to the aircraft systems on and off. In those tests, the maximum energy

released by FQIS wiring inside the fuel tank when a discharge mechanism was provided

(by placing conductive debris across FQIS fuel probe terminals) was 0.125 mJ. This is

significantly less than the 0.5 to 500 mJ that would have been needed to ignite the Jet A

fuel/air vapor in TWA flight 800ís CWT and only half of the generally accepted MIE level

for Jet A fuel (0.25 mJ). Therefore, the Safety Board concludes that it is unlikely that EMI

from aircraft system wiring played a role in igniting the fuel/air vapor in TWA flight 800ís

CWT.

603 One watt was required to induce a spark during laboratory tests.
Analysis 281 Aircraft Accident Report

2.3.2.1.2 Short Circuiting 

2.3.2.1.2.1 Possible Short-Circuit Energy Transfer Mechanisms

Excess voltage from a short circuit can be transferred from wires carrying higher

voltage to wires carrying lower voltage if the wires are near each other. Wires carrying

high voltage can be placed near FQIS wires both in common wire bundles and at

connectors in various locations. The investigation found that Boeing design specifications

permit FQIS wiring to be bundled with, or routed next to, higher-voltage airplane system

wires, some carrying as much as 350 volts.604 In addition, Safety Board examinations of

other transport-category airplanes, including several 747s, found that high-voltage wires

can be corouted with FQIS wires in ways that do not comply with the manufacturerís

production illustrations (PI) and/or are not consistent with the guidance in Boeingís

Standard Wiring Practices Manual (SWPM). These noncompliant wiring configurations

can be installed at the time of manufacture or during subsequent maintenance activities.

The CWT FQIS wiring in the 747 terminates at common connectors with the fuel

quantity indicator, the totalizer gauge, the airborne integrated data system (AIDS), the

volumetric shutoff (VSO), and the left wing refueling station; each of these connectors

also includes terminating wires from circuits that carry 115 volts.605 (See figure 46.) Short

circuits in connectors have caused incidents in the past. For example, an October 22, 1995,

incident involving roll/yaw oscillations on a 737 near Manchester, England, was attributed

to a short circuit in the rudder circuitry of the incident airplane that resulted from lavatory

fluid bridging the pins within an electrical conductor. 

A short circuit can occur if the wireís internal conductors are accessible/exposed

and there is either (1) direct contact between the conductors or (2) a bridge between the

conductors, created by contaminants, such as metal shavings or fluid.606 Conductors inside

electrical connectors are, of necessity, always accessible (because a connector is designed

to join wires in a continuous electrical path), and conductors outside connectors can

become exposed as a result of cracking or other damage to wire insulation. The

investigation found that damaged and/or contaminated wire insulation is widespread in

older transport-category airplanes and also exists in new airplanes. Specifically, Safety

Board examinations of wiring on 26 airplanes of varying ages (ranging from new to

604 Most airplane systems carry 115 volts, but in the Safety Boardís examinations of numerous transportcategory airplanes, lighting circuits of 192 and 350 volts were found in bundles that also contained wiring

associated with the FQIS. Even higher voltage wires exist in these airplanes (such as for strobe lights,

fluorescent lighting starting circuits, and radios), but none were found routed with FQIS wires.

605 Although the Safety Board is not aware of FQIS wires sharing connectors with wires carrying more

than 115 volts, Boeing specifications also permit FQIS wiring to be mixed in common connectors with

airplane system wires carrying up to 350 volts. 

606 Testing conducted for the Safety Board by Lectromec demonstrated that current peaks of more than

100 amperes (amp), which released 2 to 4 J of energy (of which about 400 mJ could be transferred to a

victim wire), could be produced on wires protected by 10-amp circuit breakers when short circuits were

created by dripping either water or lavatory fluid onto damaged wires. (For more information about this

testing, see section 1.16.6.2.)
Components:1. Depicts seven capacitive fuel probes wired in parallel.2. CWT Compensator3. T347 Terminal Strip4. Splice located behind Flight Engineer panel5. Resistor6. Fuel Flow #4 wire bundle route to forward panels, as shown in Production Illustration.Notes:A. Wires shown parallel are routed together or in parallel bundle.B. Component locations and wire relationships are approximate and shown for relation (not to scale).C. Lighter areas depict where wires were not recovered.D. Bold line depicts HI-Z circuit and circuits directly connected.

Parallel

CVR Wires

Pressure

Refueling

Panel

4.

6.

CWT

GAGE

TOTALIZER

FUEL

FLOW

VSO

CMPTR

FUEL

FLOW

CMPTR

Left

Wing

Right

Wing

Wiring Behind

Galley C,

Beneath Floor

AIDS

UNIT

3.

5.

1. 2.

To Engine #4

Figure 46. Diagram of the center wing tank fuel quantity indication system wiring on the 747 and its common con
Analysis 283 Aircraft Accident Report

28 years old) revealed that all of the older airplanes exhibited numerous examples of

mechanically damaged, chafed, cracked, and contaminated wires. The examinations also

found sharp-edged metal drill shavings (which can damage wire insulation), fluid stains,

and other potentially hazardous material in or near the aircraft wiring of old and new

airplanes.

The Safety Board is aware of several incidents that resulted from short circuits

apparently created by contaminants.607 In addition, Board-sponsored testing at Lectromec

showed that contaminants, such as water, lavatory fluid, and metal shavings, can create a

mechanism that can allow energy to inadvertently transfer between wires with cracked or

damaged insulation. The tests demonstrated that fluids could migrate between wires with

cracked or damaged insulation, reaching copper conductors at the sites of existing

insulation damage, and become a mechanism for the intermittent transfer of short bursts of

energy (referred to by Lectromec as ìflashingî or ìscintillationsî) for as long as

25 minutes without tripping circuit breakers. Further, the tests revealed that metal

shavings could cut through insulation and unintentionally transfer energy from one wire to

another wire in a bundle (or at a connector) when the shavings contacted two bare

conductors. 

Finally, in addition to the possibility of a short circuit to the CWTís FQIS wiring,

there are several ways in which a short circuit to FQIS wires associated with fuel tanks

other than the CWT could transfer energy to the CWT. First, in the totalizer gauge on the

accident airplane, the wires attached to the connector pins for the right wing main tank

FQIS and the CWT FQIS had been improperly soldered together and had subsequently

cracked apart. During examination of the gauge at Honeywell, it was determined that

electrical energy would cross the crack in the solder between the connector pins when

slightly more than 270 volts (less than is used in lighting circuits) was applied to one of

the pins.608 Second, the routing of the right wing FQIS wires to a connection in the CWT

(at terminal strip T347)609 created another possible mechanism by which a short circuit

from a higher-voltage wire in the right wing FQIS wiring could have carried excessive

voltage to the CWT. A third way in which excess energy from another tankís FQIS wiring

could enter the CWT was demonstrated during postaccident evaluation of wiring paths on

a retired 747; an electrical signal that was placed on the CWT FQIS wiring was also

detected on wiring from the left wing FQIS, indicating that the two sets of wiring were not

electrically isolated. Thus, a short circuit from a higher-voltage wire to left wing FQIS

wiring could result in excess energy being transferred to the CWT. 

2.3.2.1.2.2 Adequacy of Existing Circuit Protection Systems

As previously discussed, short circuits can transfer energy from higher-voltage

wires to lower-voltage FQIS wires and become a potential fuel tank ignition hazard. More

607 For more information about these incidents, see section 1.18.2.3.

608 Although no 270-volt circuits were identified in the wings of any airplanes examined, the PIs for the

accident airplane showed right wing FQIS wires were routed through the fuselage in bundles that contained

350-volt lighting circuits.

609 This same terminal strip installation was found to have sulfide deposits in two other 747s.
Analysis 284 Aircraft Accident Report

commonly, short circuits can cause sparks that can ignite lint or other flammable materials

and cause secondary damage to other aircraft system wires. To minimize or eliminate the

damage caused by short circuits, wiring circuits can be protected by one or more of the

following methods: (1) circuit breakers (either the standard, thermally activated type or

the recently developed processor-based arc-fault circuit breakers [AFCB]); (2) electrical

resistors; (3) shielding and/or physical separation of wires; and (4) transient suppression

devices (TSD), such as surge protection systems. The Safety Board evaluated each of

these methods to determine whether they could provide adequate protection to FQIS

wiring circuits by preventing short circuits to those wires from becoming potential fuel

tank ignition hazards.

2.3.2.1.2.2.1 Circuit Breakers

The design of thermally activated circuit breakers permits higher-than-desired

current to pass through the wire during normal operation before sufficient heat is

generated to interrupt the circuit (trip the circuit breaker). The level of heat required to trip

the circuit breaker is a function of not only amperage and voltage, but also the duration of

the event. Thus, higher-than-desired current may flow through the wire, intermittently, for

long periods of time without tripping the circuit breaker. As previously mentioned, testing

conducted for the Safety Board by Lectromec demonstrated that current peaks of more

than 100 amps, which released 2 to 4 J of energy610 (of which about 400 mJ could be

transferred to a victim wire), could be produced on wires protected by 10-amp circuit

breakers when short circuits were created by dripping either water or lavatory fluid onto

damaged wires. Further, intermittent short circuits existed for as long as 25 minutes

without tripping circuit breakers. Therefore, because the energy levels transmitted during

these types of electrical activity are high enough to ignite flammable fuel/air vapor in a

fuel tank, thermally activated circuit breakers cannot be relied upon to prevent short

circuits to FQIS wires from becoming potential fuel tank ignition hazards.

The Safety Board is aware that the FAA and industry have been working to

develop processor-based AFCBs that would trip the circuit breaker within the first

milliseconds of a higher-than-desired energy transfer. Such protection systems would

represent an improvement over thermally activated circuit breakers because they would

minimize the amount of time that current would be permitted to flow through a short

circuit (such as occurs in arc-tracking) and, therefore, also minimize the amount of fire

and system damage following a short-circuit event. Therefore, the Board supports the

continuation of efforts to develop this technology and will follow those efforts with

interest. However, the Board remains concerned that even the short duration of time that

energy is permitted to transfer by such systems might still allow the transfer of enough

energy to FQIS wiring to ignite flammable fuel/air vapor in a fuel tank. Therefore, it is

unclear whether processor-based AFCBs could adequately prevent short circuits to FQIS

wiring from becoming potential fuel tank ignition hazards.

610 On one occasion, 90 J were released during a 40-millisecond succession of flashes. 
Analysis 285 Aircraft Accident Report

2.3.2.1.2.2.2 Electrical Resistors

A resistor limits the amount of current (amperage), and thus power, that can pass

through it. According to Boeing, protective resistors in the components connected to

747 CWT FQIS wiring (fuel quantity gauges, the VSO unit, AIDs, and fuel quantity

totalizer gauge) limit the amperage from an internal short circuit to about 1 milliamp.

Therefore, although each of those components includes a 115-volt circuit, the amount of

energy transferred to the CWT FQIS wiring as a result of an internal short circuit would be

too small to ignite the fuel/air vapor in the CWT. However, as previously discussed, there

are many ways in which a short circuit outside of these components could transfer energy

to FQIS wiring. Because resistors are designed to limit current with a known voltage, they

are not appropriate devices to prevent short circuits to FQIS wiring from a variety of

possible voltage sources. 

2.3.2.1.2.2.3 Shielding and Physical Separation of Wires

As discussed in the Safety Boardís April 7, 1998, safety recommendation letter to

the FAA, shielding and physical separation of FQIS wires from other wires can be an

effective means of preventing the transfer of potentially hazardous levels of energy to

FQIS wires from higher-voltage wiring. In that letter, the Board issued Safety

Recommendation A-98-38, which asked the FAA to 

Require in Boeing 747 airplanes, and in other airplanes with FQIS wire

installations that are corouted with wires that may be powered, the physical

separation and electrical shielding of FQIS wires to the maximum extent possible.

In response, the FAA initiated two notice of proposed rulemakings (NPRM) that

subsequently led to the issuance of airworthiness directives (AD) to require shielding and

separation of FQIS wiring on 747 and 737 airplanes.611 Regarding other airplane types, the

FAA indicated that advisory material being developed in connection with the fuel tank

design reviews proposed in NPRM 99-18612 would include information ìdescribing the

use of wire separation and electrical shielding to protect FQIS wiring from short

circuitsÖand other failures and malfunctions that could induce high electrical energy on

FQIS wires that enter fuel tanks.î In its February 9, 1999, response, the Safety Board

expressed concern about the adequacy of Boeingís wire separation standard (discussed

further below) and stated that, pending receipt of further details regarding this separation

standard and planned activities for other airplane models, Safety Recommendation

A-98-38 was classified ìOpenóAcceptable Response.î

Although airplane manufacturers generally provide protection for certain critical

electrical circuits, there is no FAA regulation that specifies wire separation criteria or

identifies which circuits must be protected. In their evaluation of the adequacy of the wire

611 The NPRMs proposed the installation of TSDs as an optional or additional requirement. Although the

final ADs did not include TSDs as an alternative, the FAA indicated in correspondence related to Safety

Recommendation A-98-39 that it will nonetheless permit TSDs as an alternate means of compliance.

612 For more information about this NPRM, see section 1.18.3.4.3.
Analysis 286 Aircraft Accident Report

separation that would be required by the two ADs, Safety Board investigators reviewed

the general wire separation standards and practices of several airplane manufacturers and

found that these standards are not uniform. For example, Douglas Aircraft Company

specified that wiring for certain systems (including FQIS and other fuel system wiring,

fire warning system wiring, primary generator feeder cables, and electro-explosive

devices) must be separated by at least 3 inches from other electrical wiring. In contrast,

Boeing specifications do not require protection for some of the systems specified by

Douglas (such as the FQIS), and, for those systems that are designated as protected, the

required separation distance is only 1/4 inch in pressurized areas and 1/2 inch in

unpressurized areas.

The potential for short circuits to damage nearby wiring (more than 1 1/2 inches

away) has been documented in Safety Board investigations of numerous accidents and

incidents.613 The Safety Board concludes that existing standards for wire separation may

not provide adequate protection against damage from short circuits. Therefore, the Safety

Board believes that the FAA should review the design specifications for aircraft wiring

systems of all U.S.-certified aircraft and (1) identify which systems are critical to safety

and (2) require revisions, as necessary, to ensure that adequate separation is provided for

the wiring related to those critical systems. 

2.3.2.1.2.2.4 Transient Suppression Devices

TSDs limit the amount of energy or current that can pass through them to a

predetermined amount. In its April 7, 1998, safety recommendation letter to the FAA, the

Safety Board indicated that, because of the number of potential latent ignition sources in

fuel tanks, the many means by which energy can be introduced into FQIS wires, and the

difficulty of achieving complete wire separation and shielding in in-service airplanes,

TSDs (referred to in that letter as ìsurge protection systemsî) may be the most effective

method of preventing FQIS wiring from becoming an ignition source. The Board noted

that if such devices were installed where FQIS wires enter fuel tanks, they could provide

added protection against excessive power surges in the FQIS system, regardless of origin.

Accordingly, the Board issued Safety Recommendation A-98-39, which asked the FAA to

Require, in all applicable transport airplane fuel tanks, surge protection systems to

prevent electrical power surges from entering fuel tanks through FQIS wires.

In its June 10, 1998, response, the FAA indicated that it had issued NPRMs

proposing to require, on 747s and 737s, installation of TSDs and/or installation of wire

shielding and separation of the electrical wiring of the FQIS located outside of the fuel

tanks. In its February 9, 1999, response, the Safety Board stated that wire shielding and

613 For example, during postincident inspection of a shorted wire bundle after a January 9, 1998, incident

aboard a United Airlines 767 en route from Zurich, Switzerland, to Washington Dulles International Airport

in which the flight crew made a precautionary landing at London Heathrow International Airport after

experiencing multiple display errors and tripped circuit breakers, power was inadvertently reapplied and

arcing resumed, but it did not trip any additional circuit breakers. However, this event damaged dozens of

wires, some more than 1 1/2 inches away, and also burned printed circuit tracks from cards located in

another section of the electronics compartment.
Analysis 287 Aircraft Accident Report

separation alone could not protect against the entry of power from sources attached to

FQIS wires and noted that FQIS wires share common connections with some

higher-voltage circuits. The Board urged the FAA to reconsider its response to this safety

recommendation. On the basis of the FAAís decision not to require TSDs, the Board

classified A-98-39 ìOpenóUnacceptable Response.î

In its November 3, 1999, and May 17, 2000, letters, the FAA indicated that on the

basis of comments received on the NPRM, it had initially determined that TSD technology

was not sufficiently developed for application to in-service airplanes and that wire

separation and shielding provided a better design solution for protection of FQIS wiring.

The FAA also stated that the requirement for separation and shielding of electrical circuits

would apply to all portions of the circuit (including those portions within connectors and

components).

Regarding the reliability of TSDs, the FAA indicated in its November 3, 1999,

letter that it had been particularly concerned about comments that it had received on the

NPRM regarding the potential for latent failure modes and the introduction of error in fuel

quantity indications. However, the FAA further stated that several manufacturers had

subsequently expressed interest in developing TSD technology as a possible alternate

means of compliance with the ADs and that approval of TSDs would be considered. The

FAA indicated that it had started determining appropriate requirements for such approval.

The FAA further indicated that it did not plan to propose a requirement for both circuit

separation and shielding and the installation of TSDs and that it believed either design,

executed properly, would adequately ensure that high-energy signals are not introduced on

wiring inside fuel tanks.

In its May 17, 2000, letter to the Safety Board, the FAA stated that the

requirements for approval of a TSD design will include demonstrating that the device will

limit the transfer of energy and power through the wires that enter a fuel tank to

intrinsically safe levels (limited to less than spark energy and filament heating limits that

could create an ignition spark or a hot surface ignition source in a fuel tank). According to

the FAA,614 in connection with approving TSDs as an alternate means of compliance, it

will be assumed that failure of the TSD will result in catastrophic consequences.

Therefore, it appears that for a TSD design to receive approval, it would have to be shown

that the possibility of any such failure is extremely improbable, as defined in Advisory

Circular (AC) 1309-1A. 

As will be discussed in section 2.4.2, the FAA has initiated rulemaking to require

minimizing fuel tank flammability in newly designed airplanes and is evaluating methods

to reduce flammability in existing designs. If flammability in existing and future transportcategory airplanes were eliminated or reduced (to the maximum extent technologically

feasible), the Safety Board would agree with the FAA that either TSDs or separation and

shielding, if executed properly, will provide adequate protection against introduction of

excess energy to wiring inside fuel tanks.615 Therefore, pending the timely completion of

614 See Issue Paper P-1, ìFQIS Transient Suppression Device/Unit for Installation on Boeing

Model 747-100, -200, SP and SR Series Airplanes,î dated January 13, 2000.
Analysis 288 Aircraft Accident Report

the FAAís actions to eliminate or reduce (to the maximum extent technologically feasible)

flammability, Safety Recommendations A-98-38 and -39 are classified ìOpenó

Acceptable Response.î

2.3.2.1.2.3 Factors Suggesting the Likelihood that a Short-Circuit Event 

Occurred on TWA Flight 800

Much of the insulation on the wiring recovered from the accident airplane was

cracked or otherwise damaged, often exposing the inner conductor. When powered, such

damaged wires would be vulnerable to short-circuiting. Although some of the damage to

the accident airplaneís wiring insulation probably occurred as a result of the accident or

search and recovery operations, the degraded condition of wiring insulation found during

inspections of other transport-category airplanes of about the same age as the accident

airplane (and representing numerous different operators) suggests that at least some of the

damage to the wiring insulation on the accident airplane very likely existed before the

accident. Given what was found during the inspections of other airplanes, it is also likely

that metal shavings and other contaminants were interspersed with the wiring system on

the accident airplane before the accident. 

Evidence of arcing was found on generator cables routed with wires in the leading

edge of the right wing, near the wing root. Although this arcing damage might have been

caused by the breaking of the forward wing spar and subsequent fuel fire, it is possible that

it could also have been caused before the explosion. Because this wire bundle included

wires leading to the right main wing tank (No. 4) fuel flow gauge and right wing FQIS

wiring that would have been routed to a connection in the CWT at terminal strip T347, a

short circuit in this bundle could have carried excessive energy into the CWT FQIS. 

In addition, two non-FQIS wires at body station (STA) 955, which would have

been corouted in the same raceway as CWT FQIS wiring, were found with possible arcing

damage.616 (Although the FQIS wiring recovered from this area did not contain evidence

of arcing, it should be noted that some of the FQIS wiring from this area was not

recovered.) These wires were located near structural repairs from a burst potable water

tank and numerous other floor repairs. These repairs could have disturbed nearby wires,

cracking or otherwise damaging the wire insulation, and could also have generated metal

shavings. In fact, metal drill shavings were found adhered to fragments of a floor beam

from STA 920, within 2 inches of where the CWT FQIS wiring would have been routed.

This area was also near galley C, which was the site of numerous reported leaks in the

2 weeks preceding the accident. Leakage from this area could have dripped onto electrical

615 The Safety Board recognizes that some FQIS technologies currently in development do not rely on

wiring inside fuel tanks. Instead, the sensors for such systems utilize fiber-optics, sonics, and other

underlying mechanisms. No additional protection (TSDs or separation and shielding of wires) would be

required to prevent energy from entering fuel tanks so equipped.

616 Copper strands on these two wires had melted and rehardened with molten/resolidified ends.

Although it is possible that this damage could have been caused by the postexplosion fire, investigators

considered it unlikely that only two wires from this raceway would have sustained such damage if the entire

area had been exposed to fire.
Analysis 289 Aircraft Accident Report

wiring located immediately beneath the galley floor and caused a short circuit that affected

the CWT FQIS wiring. 

Repairs to the area around the upper deck flight attendant lighting panel could also

have created conditions conducive to short-circuiting. A lighting wire and pin in that panel

had been repaired on June 20, 1996, about a month before the accident. Although no

evidence of arcing was found on the repaired wire, during the repair other wires bundled

with it might well have been moved. The repaired wire was part of a bundle that branched

off from a larger bundle that contained CWT and left wing FQIS617 wires that led to the

upper deck AIDS unit and also contained high-voltage wiring for lighting; thus,

manipulation of wires during the repair could have resulted in movement and cracking of

these wires. In addition, there was evidence of extensive structural repairs in this area, and

the cabin interior had been altered, both of which could have disturbed these wires and

introduced metal shavings, possibly damaging the wire insulation. Further, condensation,

which is common in transport-category airplanes,618 could have provided a mechanism for

short-circuiting of such damaged wires when powered. Finally, in addition to being

bundled with FQIS wires, the lighting wires were also bundled with CVR wires and No. 4

fuel flow wires along some portions of their path. Therefore, a short circuit involving

these lighting wires could also explain the electrical anomalies indicated on the CVR

recording and the No. 4 fuel flow indicator.

Although no evidence of arcing was found in any of the components connected to

the CWT FQIS (the fuel quantity indicator, totalizer gauge, AIDS, VSO unit, and left wing

refueling station), investigators considered the possibility that a short circuit in one of

those components could have been a source of excess voltage transferred to the CWT

FQIS wiring. The interior of each of these components contained numerous complex

wiring and circuit assemblies that could have obscured evidence of a short circuit. Further,

it is also possible that a short circuit at lower power or through moisture could occur

without leaving evidence of arcing.619 Therefore, there were several possible locations at

which a short circuit of higher-voltage wiring could have affected the CWT FQIS wires in

the accident airplane. 

Further, as noted previously, there are several indications that possible anomalous

electrical events occurred in the airplane just before the explosion. First, the captainís

CVR channel recording has two ìdropoutsî of background power harmonics, indicating

some type of electrical anomaly, less than a second before the CVR lost power.620 Second,

captainís comments about a ìcrazyî No. 4 (right outboard engine) fuel flow indicator were

617 As previously noted, FQIS wiring from wing tanks is not electrically isolated from the CWT FQIS

wires and, therefore, energy resulting from a short circuit to any of those FQIS wires could potentially reach

the CWT. 

618 The 747 design includes drip shields between the upper deck and cockpit to protect critical elements

(including the cockpit wiring) from condensation. 

619 For more information about arcing, see section 1.16.6.2.

620 The continuation of the 400- and 800-Hz signals through these two periods indicate that the CVR was

functioning and that the dropouts were not the result of damage to the recording. Safety Board testing found

that power interruptions on adjacent circuits could result in acoustical indications on the CVR.
Analysis 290 Aircraft Accident Report

recorded on the CVR about 2 1/2 minutes before it lost power, which also suggests that

some type of electrical anomaly occurred that affected that wiring. And third, the

recovered CWT fuel quantity gauge from the cockpit displayed a reading of 640 pounds,

which does not agree with the quantity recorded by the ground refueler (300 pounds).621

Safety Board testing showed that applying power to a wire leading to the fuel quantity

gauge can cause the digital display to change by several hundred pounds in less time than

is required to trip the circuit breaker. This suggests that an electrical anomaly might have

affected the reading of the cockpit gauge (accounting for the discrepancy). 

These electrical anomalies were not necessarily related to the same event.

However, it is possible that one or more of these anomalies were a manifestation of an

electrical event that resulted in excess voltage being transferred to the CWT FQIS wiring.

On the basis of this and the other evidence previously discussed, the Safety Board

concludes that a short circuit producing excess voltage that was transferred to the CWT

FQIS wiring is the most likely source of ignition energy for the TWA flight 800 CWT

explosion.

2.3.2.2 Release of Ignition Energy Inside the Center Wing Fuel Tank

2.3.2.2.1 Arcing

Exposed conductors on FQIS wiring (caused by either mechanical damage or

cold-flow)622 within a fuel tank could provide a mechanism that would lead to arcing

inside the tank, which in turn could ignite the flammable fuel/air vapor.623 (This was one

of the suspected ignition mechanisms for the 1990 Philippine Airlines 737 CWT

explosion and the 1972 Navy C-130 fuel tank explosion.) Very little (about 4 feet) of the

CWT FQIS wiring from the accident airplane was recovered, and, therefore, the degree to

which the wiring in the tank might have been damaged before the accident could not be

assessed. However, investigators found preaccident damage, including exposed

conductors,624 on some of the recovered FQIS wiring from inside TWA flight 800ís wing

tanks, and damaged FQIS wiring was found inside the CWTs of several of the other 747

airplanes examined by the Safety Board.625 In addition, the presence of a conductive

material, such as metal drill shavings or safety wire (both of which were found in the fuel

621 There is evidence that the cockpit gauge indicated 300 pounds while the airplane was on the ground;

the figure recorded by the ground refueler (300 pounds) would have been taken from a ìrepeaterî gauge at

the refueling station, which registers the same quantity as the cockpit gauge. The difference between this and

the 640-pound indication at the time of the accident exceeded the accuracy tolerance allowed by the

manufacturer and the error recorded during certification tests. Further, the 640-pound indication was found

in both the electrical and mechanical gauge mechanisms, which also suggests that the indication may not

have been attributable to an error in the gauge. 

622 Cold-flow refers to migration or displacement of wire insulation.

623 Investigators also theorized that arcing might occur if a loosely connected FQIS probe intermittently

contacted structure inside the fuel tank. However, no loose fuel probes were found during FAA-required

postaccident inspections of 747s for loose fuel probes, and there was no evidence in the recovered wreckage

from the accident airplane to support this theory. 

624 Some of the exposed core conductors resulted from cold-flow of the polytetrafluoroethylene

insulation material, and others apparently resulted from improper maintenance. Sulfide deposits were found

on some exposed conductors, indicating that the damage existed before the accident.
Analysis 291 Aircraft Accident Report

tanks of other airplanes), could have provided a mechanism that would lead to arcing of

FQIS components. Although no clear evidence of arcing was found inside TWA flight

800ís CWT, fire damage along the route of the FQIS wiring was severe enough that it

likely would have obscured any such evidence.

2.3.2.2.2 Resistance Heating

Another potential source of ignition energy is resistance heating, which could have

resulted from a thin filament626 being heated through contact with a wire, probe, or

compensator exposed to excess voltage.627 Although no clear evidence of a filament

ignition was found inside TWA flight 800ís CWT, such evidence could also have been

physically lost or obscured by fire damage. 

2.3.2.2.3 Silver-Sulfide Deposits on Fuel Quantity Indication System Wiring 

and Components

Research and testing conducted during this investigation found that silver-coated

copper parts inside fuel tanks, such as those used in the FQIS, can develop silver-sulfide

deposits628 that are semiconductive and, therefore, can reduce the resistance between

electrical connections and permit arcing. Such deposits can become a potential ignition

mechanism inside a fuel tank.629

Silver-sulfide deposits were found on FQIS probes, compensators, and wiring

from the accident airplane. Similar deposits have been discovered on FQIS parts from

other airplanes, both military and civilian (including other 747s). Research by the Safety

Board, the FAA, and the AFRL has found that deposits accumulate over time as a result of

the partís exposure to jet fuel and fuel vapors, which contain sulfur. (It should be noted

that the FQIS design for the 747 classic includes exposed conductors at terminal blocks.)

In addition, laboratory research by the University of Dayton Research Institute conducted

as a result of this investigation found that increased temperatures enhanced the formation

of silver-sulfide deposits and that fuel washing could slow the accumulation of these

deposits. This finding is significant with regard to CWTs because they are less frequently

625 This damage was discussed in the Safety Boardís April 7, 1998, safety recommendation letter to the

FAA. 

626 A filament is a slender wire, such as that found inside an electric light bulb.

627 For example, testing by Boeing indicated that steel wool can be ignited by as little as 32 milliamps

from a 25-volt 400-cycle alternating current source. (The Boeing data did not include variables required to

calculate energy [J] from the voltage value that was provided.)

628 These deposits are sometimes also referred to as ìcopper-sulfide deposits,î ìsulfides,î or

ìsulfidation.î

629 Ignition of vapors in an AFRL trainer fuel probe in 1990 during electrical bench tests (when test

voltages higher than normal fuel system voltages were applied) was attributed by the AFRL to sulfide

deposits. Subsequent AFRL testing for the Safety Board indicated that the deposits could break down and

result in arcing when exposed to a 170-volt pulse. Also, during the investigation of the TWA flight 800

accident, the Board became aware that in 1988, Boeing found evidence of silver-sulfide deposits at the

location of arcing in 18 fuel pump motors. In addition, during FAA-sponsored testing at Arizona State

University in connection with this investigation, application of direct current voltage to silver-sulfide

deposits created in a laboratory ignited Jet A fuel vapors. 
Analysis 292 Aircraft Accident Report

filled with fuel to the top than wing tanks, and, therefore, the terminal block at the top of

the CWT fuel probes is less often covered by fuel (which would subject it to fuel washing)

than would be the case in a wing tank.630

The AFRL concluded that the accumulation of silver-sulfide deposits ìis most

likely the result of a long-term degradation or corrosion processÖ[and that] as the probes

age, more probe failures [calibration errors] can be expected.î However, the problem is

not solely limited to older airplanes; silver-sulfide deposits that were significant enough to

cause FQIS anomalies631 were found in a 757 airplane with only 750 hours.632

The existence of silver-sulfide deposits on FQIS components from both the

accident airplane and numerous other airplanes raises the possibility that sulfide deposits

could have provided an energy release mechanism at the point of ignition in the CWT. The

accident airplaneís recent maintenance history, which indicates FQIS anomalies, is

consistent with silver-sulfide deposits interfering with FQIS operation. Accordingly,

sulfide deposits were considered as a possible ignition mechanism for the TWA flight 800

CWT explosion.

2.3.2.2.3.1 Elimination of Silver-Sulfide Deposits Inside Fuel Tanks

In response to AFRLís findings from the 1990 vapor ignition incident,

BFGoodrich eliminated the use of silver-plated components in the FQIS and began using

nickel-plated wire, gold-plated ring connectors, and sealant. Those improved components

have been used in military airplanes since about 1993. According to BFGoodrich, since

that time, there has been a large reduction in the FQIS anomalies that had been associated

with silver-sulfide deposits (such as FQIS inaccuracies). Although a Boeing 1991

engineering report indicated that silver should never contact sulfur-containing liquids

because of the susceptibility to sulfidation, and Boeing uses nickel-plated (instead of

silver-plated) wiring in its newly manufactured 777 and 737-NG airplanes, Boeing

indicated in a December 7, 1999, letter to the Safety Board that it does not recommend

replacing silver-plated FQIS components in existing airplanes.

In an April 7, 1998, letter to the FAA, the Safety Board expressed concern that

sulfide deposits on FQIS wires could become an ignition source in fuel tanks and issued

Safety Recommendation A-98-37, which asked the FAA to

Require research into copper-sulfide deposits on FQIS parts in fuel tanks to

determine the levels of deposits that may be hazardous, how to inspect and clean

the deposits, and when to replace the components.

630 Only two of the many FQIS components from other airplanes examined during this investigation that

contained silver-sulfide deposits had been removed from the airplane because of FQIS anomalies; both were

terminal strips from the top of 747 CWTs. 

631 Research by BFGoodrich, an FQIS manufacturer, conducted after the AFRL vapor ignition incident

in 1990 found that silver-sulfide deposits could permit leakage of electrical currents and, therefore, cause

FQIS anomalies.

632 For more information about sulfide deposit tests, see section 1.16.6.9.
Analysis 293 Aircraft Accident Report

In response to this recommendation, the FAA convened a team of specialists from

Government, academia, and industry to conduct research. According to the FAAís May 17,

2000, response to the Safety Board, the team has studied the properties of sulfide deposits

created in a laboratory environment but has not found indications of significant levels of

deposits occurring in in-service airplanes for comparison. The FAA indicated that the

research program had been extended from its original 12-month timeframe so that the

team could complete its evaluation. 

The FAA indicated in NPRM 99-18 that it anticipates that the proposed fuel tank

safety review will identify critical areas of the fuel tank system that will require

maintenance actions to account for the effects of aging, wear, corrosion, and possible

contamination of the fuel system. Specifically, the FAA noted that it might be necessary to

provide maintenance instructions to identify and eliminate silver-sulfide deposits.633

The Safety Board concludes that silver-sulfide deposits on FQIS components

inside fuel tanks pose a risk for ignition of flammable fuel/air vapor. Although precluding

operations with flammable fuel/air vapor (as recommended in Safety Recommendation

A-96-174) would eliminate this risk, flammable fuel/air vapors will not likely be

eliminated in all fuel tanks or during all phases of flight. Therefore, although the Board is

pleased that the FAA is continuing to support research into sulfide deposits, the Safety

Board believes that the FAA should require the development and implementation of

corrective actions to eliminate the ignition risk posed by silver-sulfide deposits on FQIS

components inside fuel tanks. The Safety Board classifies Safety Recommendation

A-98-37 ìClosedóAcceptable ActionóSuperseded.î

2.3.2.2.4 Ignition Location Inside the Center Wing Fuel Tank

The Safety Board contracted with two research laboratories, Sandia National

Laboratory and Christian Michelsen Research, to develop computer modeling in an

attempt to determine potential ignition locations. However, because of considerable

uncertainties in some aspects of the methodology, the results of that modeling could not be

used to determine the most likely ignition location.634

Nonetheless, investigators examined all of the recovered CWT FQIS components,

which included portions of all seven fuel probes, one complete terminal block and one

partial terminal block, and the compensator.635 None of the recovered probes or terminal

blocks exhibited any noteworthy signs of damage. However, several plastic parts inside

the compensatorís innermost tube were found burned, with an apparent upward-flowing

burn pattern, which investigators hypothesized could indicate that a fire initiated inside the

633 In NPRM 98-NM-163-AD, the FAA proposed that silver-plated copper FQIS wiring in 747 CWTs be

replaced with nickel-plated copper FQIS wiring within 20 years of the airplaneís manufacture. Although the

FAA acknowledged that there was little technical basis for this time limit, and the proposed replacement

requirement was not included in the final rule (AD 98-08-02), the Safety Board is concerned that this

proposal indicates a belief on the part of the FAA that silver-sulfide deposits do not pose a significant risk

for airplanes less than 20 years old. As discussed in section 2.3.2.2.3, potentially hazardous deposits can

accumulate in younger airplanes as well as on older airplanes, and such deposits can pose a significant risk. 

634 For more information about the computer modeling, see section 1.16.5.5.
Analysis 294 Aircraft Accident Report

compensator. Similar burn patterns636 were observed on the compensator believed by the

FAA to be the ignition source for the surge tank fire in the 747 that experienced a fuel tank

explosion in May 1976 near Madrid, Spain.637 (Although the Safety Board discounted the

compensator as an ignition source in its October 1978 report of the Madrid accident, a

different conclusion might have been warranted638 given what is now known about

sulfides and other ignition-related phenomenon.) However, there was insufficient

evidence to conclude that the damage to the CWT compensator occurred before the

explosion, and, therefore, no determination could be reached regarding the likelihood that

the compensator was the ignition location. 

2.3.2.3 Possible Ignition Scenario for TWA Flight 800 Explosion

Therefore, the Safety Board concludes that the ignition energy for the CWT

explosion most likely entered the CWT through the FQIS wiring, and, although it is

possible that the release of ignition energy inside the CWT was facilitated by the existence

of silver-sulfide deposits on an FQIS component, neither the energy release mechanism

nor the location of the ignition inside the CWT could be determined from the available

evidence.

2.4 Fuel Tank Design and Certification Standards

2.4.1 Preclusion of Ignition Sources in Fuel Tanks

As previously discussed, a fuel tank explosion requires both a flammable fuel/air

mixture and an ignition source. For purposes of certification, the FAA and

transport-category airplane manufacturers have historically assumed that a flammable

fuel/air mixture exists in fuel tanks at all times and have attempted to preclude fuel tank

explosions by eliminating ignition sources in fuel tanks.639 This philosophy is reflected in

NPRM 99-18, which proposes to require (1) substantiation that no ignition sources will

exist in fuel tanks and (2) development of maintenance and inspection criteria to ensure

635 Although the burned compensator did not have any marks identifying it as being from the CWT, it

was identified as such by process of elimination. Only one fuel tank other than the CWT (reserve tank No. 4)

showed any evidence of fire in the vicinity of a compensator; the compensator from this tank was positively

identified by unique characteristics. Therefore, the burned compensator apparently came from the CWT. 

636 For more information about these burn patterns, see section 1.18.1.

637 Although the ignition energy for that explosion is believed to have come from a lightning strike, the

energy would still have had to find a release mechanism inside the tank. 

638 The Safety Board explained its determination that ìfuel did not igniteî at the compensator by noting

that ì[t]ests showed that it took more voltage to spark the probe than would conceivably be induced in its

wiring [and that] microscopic examination found no evidence of sparking.î However, as was pointed out in

an undated FAA memo commenting on the Board report, the report also recognized that ìenergy levels

required to produce a spark will not necessarily damage metal or leave marksî and that ìa spark [of]

Ösufficient energy to ignite the fuelÖwill not necessarily leave physical evidence.î 

639 As stated in the FAAís January 12, 2000, Draft AC 25.981-1X, ìFuel Tank Ignition Source Prevention

Guidelines,î ì[d]ue to the difficulty in predicting fuel tank flammability, the FAA has always assumed that a

flammable fuel air mixture exists in airplane fuel tanks and has required that no ignition sources be present.î
Analysis 295 Aircraft Accident Report

that the level of safety originally designed into the fuel tank system is maintained.640 The

proposed rules would require affected type certificate and supplemental type certificate

holders to create a failure modes and effects analysis (FMEA) for all components in the

fuel tank system and use the resulting data to develop a quantitative fault tree analysis to

determine whether combinations of failures expected to occur in the life of the fleet could

cause an ignition source to exist in a fuel system. In the NPRM, the FAA indicated that

these analyses should conform to the structured requirements of 14 CFR

Sections 25.901(c) and 25.1309, which the FAA acknowledged ìhave not been

consistently applied and documented when showing that ignition sources are precluded

from transport category airplane fuel tanks.î641

The Safety Board recognizes that FMEAs and fault tree analyses provide a

structured and systematic method for identifying and evaluating the potential risk that

might be posed by failure modes and, therefore, that they can be a useful part of the

certification process. However, the Board is concerned that undue reliance is being placed

on such analyses as proof that ignition sources have been precluded. As further discussed

below, there are several reasons why such analyses may be unreliable.

Unrealistic or otherwise flawed data can be used to develop fault tree analyses.

This problem was demonstrated during this investigation by a CWT ignition fault tree

analysis that was developed by Boeing at the Safety Boardís request (but that was not

required to be and was not submitted as part of the FAA certification process). The

exposure times and failure rates appeared to be too low for some events, resulting in

overly optimistic probability data. Failure analysis specialists at NASA, who reviewed the

fault tree analysis at the Boardís request, stated the following in a November 25, 1998,

letter: 

Many of the probabilities, failure rates, and/or exposure times were much lower

than would reasonably be expected. The probability of occurrence should be

higher and/or exposure times should be longer on many of the basic

events....Based on our evaluation of the tree and the information provided by the

[Safety Board], the subject fault tree analysis quantification cannot stand up to

peer review and should not be viewed as realistic. It should be noted that the logic

of the tree could not be fully evaluated since we did not have access to the

engineering drawings and schematics of the system. However, based on previous

systems experience, we would expect the tree to be constructed differently. 

Further, Safety Board staff noted that the two formats in which the data were

presented (graphical and tabular) indicated different levels of risk for an ignition source

(which was identified as the top-level failure mode); one form indicated that the risk of a

CWT explosion was within acceptable limits (that is, it was extremely improbable, as

640 The NPRM also addresses flammability reduction in newly designed airplanes; however, it does not

address flammability in current designs.

641 These regulations require the applicant to show (using a quantitative evaluation) that no failure or

probable combination of failures will prevent continued safe flight and landing. For more details about these

requirements, see section 1.16.6.5.
Analysis 296 Aircraft Accident Report

defined in AC 1309-1A), and the other indicated that the risk of a CWT explosion was not

within acceptable limits.642

In its submission, Boeing indicated that, because the fault tree analysis was

developed only to support the accident investigation, it did not go through the ìiterative

review process with the FAA normally associated with certification of a system, nor were

the underlying FMEAs developed that normally would be the source of data to build the

[fault tree analysis].î The Safety Board recognizes that the data submitted by Boeing to

the Board in connection with this investigation might not have been submitted to, or

accepted by, the FAA during the certification process. However, it is unclear whether the

problems identified by the Board and NASA would have been avoided or corrected during

that process, since Boeing was aware that the information it was supplying to the Board

would be used as part of the TWA flight 800 investigation. In any event, the questionable

data used by Boeing in its CWT ignition fault tree analysis raises concerns that similarly

questionable data may have been used to develop other fault tree analyses that have been

submitted to and accepted by the FAA in connection with aircraft certifications. The

adequacy of fault tree analyses submitted in connection with aircraft certification was not

examined as part of the TWA flight 800 accident investigation. However, this issue will be

addressed in the Boardís upcoming safety study on aircraft certification. 

Contributing to the possibility that FMEAs and fault tree analyses will be based on

inaccurate data and, therefore, will not accurately depict all possible failure modes and

their probability of occurrence is the fact that no single source exists for reliable and

comprehensive data on component failures or malfunctions. Because the calculations in an

FMEA are based (in part) on failure rates, incomplete or inappropriate failure data can

significantly skew the results of an examination. The Safety Board is aware that service

history data maintained by manufacturers do not include data from all operators. Further,

the Board has found that the amount of data provided by the manufacturers of replacement

component parts sometimes greatly exceeds the data provided by the aircraft

manufacturers (possibly because replacement part suppliers can sell parts directly to

operators and repair facilities). Although the FAA collects a significant amount of data

about mechanical failures through its Service Difficulty Report (SDR) program, even

these data are incomplete.643 Other sources of potentially relevant data are the service

histories maintained by the military of its variants of commercial airliners and the Boardís

accident and incident investigation database; however, neither of these sources provides

complete data either. 

Even a well-constructed FMEA or fault tree analysis that incorporates the most

complete and accurate data available will be unreliable to some extent because it is not

possible to predict all failure modes or to accurately determine their probability of

occurrence. As demonstrated by the series of accidents and incidents involving

uncommanded rudder movements on 737s and the recent horizontal stabilizer trim system

failure on a McDonald Douglas MD-83, unanticipated failure modes in aircraft systems

642 For more information about Boeingís CWT ignition fault tree analysis, see section 1.16.6.5.

643 For more information about concerns regarding and actions taken to address problems with the SDR

program, see section 1.18.2.1. 
Analysis 297 Aircraft Accident Report

can and do occur. The FAA has acknowledged that ìignition sources have developed in

airplane fuel tanks due to unforeseen failure modes or factors that were not considered at

the time of original certification of the airplane.î644 Similarly, the engineering and test

evaluation board that was convened in 1999, which was in response to a Safety Board

recommendation, to evaluate the 737 rudder system identified catastrophic failure modes

that had not been identified through the FMEA and fault tree analyses submitted in 1997

in connection with certification of the 737-NG airplanes. As evidenced by the Space

Shuttle disaster, even systems that have undergone the rigorous and highly respected

NASA failure analysis reviews can experience unexpected failure modes. Similarly, the

March 28, 1979, nuclear power accident at the Three Mile Island nuclear power plant near

Harrisburg, Pennsylvania, was attributed to a series of failures that were not anticipated by

the Nuclear Regulatory Commission and, therefore, not considered during

development.645

In addition, FMEAs are sometimes intentionally constructed to exclude certain

failure modes. The FAA acknowledged in draft AC 25-981-1X that the fuel tank ignition

FMEAs proposed in NPRM 99-18 are not intended to consider all potential ignition

sources. Ignition sources that are outside the purview of the proposed FMEAs include

uncontained engine failure debris; terrorist actions, such as bombs; and the effects of

postcrash or other fires outside the tank (including the propagation of fire through the

airplane vent system). 

Finally, to the extent that the prevention of failure modes are dependent on

adequate maintenance and inspection of systems and components, the premise that

maintenance and inspection programs will prevent failures is also inherently flawed.

Many potentially hazardous conditions in aircraft electrical wiring and related fuel system

components (such as worn wiring or exposed conductors obscured by clamps or other

wires, debris inside fuel pumps, and hidden silver-sulfide deposits) are not easily detected

by visual inspection, which is currently the primary method for inspecting such

components. Other potentially hazardous conditions (such as metal shavings or other

debris) may be perceived by maintenance personnel as commonplace and, therefore,

insignificant. Consequently, necessary maintenance and/or replacement of some

componentsóupon which a favorable risk assessment may have been basedómay not

occur. 

Therefore, although FMEAs and fault tree analyses focused on eliminating

ignition sources are important and necessary parts of the design and certification process,

the Safety Board concludes that they should not be relied upon as the sole means of

demonstrating that an airplaneís fuel tank system is not likely to experience a catastrophic

failure. The Safety Board further concludes that a fuel tank design and certification

philosophy that relies solely on the elimination of all ignition sources, while accepting the

644 FAA Draft AC 25.981-1X, ìFuel Tank Ignition Source Prevention Guidelines,î dated January 12,

2000. Similarly, the FAA acknowledged in NPRM 99-18 that the TWA flight 800 accident investigation has

led to the discovery of numerous previously unforeseen failure modes that could result in fuel tank ignition

and that were not considered in connection with the 747 original certification.

645 For more information about the Three Mile Island accident, see section 1.16.6.5.1.
Analysis 298 Aircraft Accident Report

existence of fuel tank flammability, is fundamentally flawed because experience has

demonstrated that all possible ignition sources cannot be predicted and reliably

eliminated. 

The design and certification of the 737 rudder system and the MD-80 stabilizer

trim system may have been similarly flawed because both of these systems have

experienced unanticipated catastrophic failures646 as a result of failure modes that appear

to have been unforeseen. The flawed 737 rudder system design and the potentially flawed

MD-80 horizontal stabilizer design would have been precluded by a certification approach

that required a reliable independent means for overcoming or counteracting the

catastrophic effects of a failure, even if the probability of the catastrophic failure mode is

calculated to be extremely improbable. (This issue will be further addressed in the Safety

Boardís upcoming safety study on aircraft certification.) In the case of fuel tank safety,

such redundant protection could be achieved by requiring, in addition to the preclusion of

ignition sources, the preclusion of a flammable fuel/air mixture. The Safety Board

concludes that operating transport-category airplanes with flammable fuel/air mixtures in

fuel tanks presents an avoidable risk of an explosion.647

2.4.2 Preclusion of Flammable Fuel/Air Vapor in Fuel Tanks

A major reason for the flammability of the fuel/air vapor in the CWT on the 747

(and similarly designed airplanes)648 is the large amount of heat generated by the air

conditioning packs located directly below the tank, which significantly elevates the

temperatures in the pack bay. Heat from the pack bay can transfer to the CWT through the

bottom of the tank and cause temperatures to rise above the lower flammability limit.649

Therefore, the 747 (and other similarly designed airplanes) may operate a significant

portion of the time with a flammable fuel/air mixture.650 Under these circumstances, a

single ignition source could cause an explosion and destroy the airplane. 

646 In the case of the 737, the catastrophic failures were uncommanded rudder movements; in the case of

the MD-80, they were uncommanded horizontal stabilizer movements.

647 The FAA and the aviation industry have also recognized that there is a safety benefit in reducing the

degree of fuel tank flammability. For example, after several fuel tank explosions in the 1970s, the use of

JP-4 fuel in civil transport-category aircraft was discontinued in favor of Jet A fuel, which is less flammable.

Further, in NPRM 99-18, the FAA proposed regulatory changes aimed at minimizing the development of

flammable fuel/air mixtures in newly designed airplanes. However, as further discussed in section 2.4.1.2,

no regulatory action has yet been taken to address the flammability of fuel tanks in the existing fleet and in

current designs.

648 The air conditioning packs are also located under the CWT in several Boeing and Airbus airplanes;

however, in some Airbus, airplanes the pack bays are ventilated. In contrast, in the DC-10, MD-11, and

L 1011, the packs are located in the nose; in the MD-80, packs are located in the tail.

649 Further, as shown in the Safety Boardís July 1997 flight tests at JFK, ullage temperatures are highest

at the bottom and center of the CWT (nearest the heat source) and become cooler in the regions around the

side and top of the CWT (where the heat leaves the tank).
Analysis 299 Aircraft Accident Report

In its December 13, 1996, letter to the FAA, the Safety Board stated that the

existence of a flammable fuel/air mixture in transport-category airplane fuel tanks was

inconsistent with the basic tenet of transport-aircraft design that no single-point failure

should prevent continued safe flight and landing. Because of its concerns, the Board

issued Safety Recommendations A-96-174 and -175, which asked the FAA to

Require the development and implementation of design or operational changes

that will preclude the operation of transport-category airplanes with explosive

fuel/air mixtures in the fuel tanks:

(a) Significant consideration should be given to the development of airplane

design modifications, such as nitrogen-inerting systems and the addition of

insulation between heat-generating equipment and fuel tanks. Appropriate

modifications should apply to newly certificated airplanes and, where feasible, to

existing airplanes. (A-96-174)

(b) Pending implementation of design modifications, require modifications in

operational procedures to reduce the potential for explosive fuel/air mixtures in

the fuel tanks of transport-category aircraft. In the 747, consideration should be

given to refueling the CWT before flight whenever possible from cooler ground

fuel tanks, proper monitoring and management of the CWT fuel temperature, and

maintaining an appropriate minimum fuel quantity in the CWT. (A-96-175) 

Initially, the FAA did not agree with the Safety Board that such changes were

necessary to address fuel tank flammability, and in its early responses relating to Safety

Recommendations A-96-174 and -175 (dated February 18, 1997, and June 28, 1997), it

indicated that it was focusing its efforts on eliminating ignition sources. However, in

October 1999, the FAA stated in NPRM 99-18 that it was undertaking a ìparallel effort to

address the threat of fuel tank explosions by eliminating or significantly reducing the

presence of explosive fuel/air mixtures within the fuel tanks in new type designs, in

production, and the existing fleet of transport airplanes.î The FAA further acknowledged

in the NPRM that the placement of air conditioning packs under fuel tanks ìwithout

provisions to reduce transfer of heat from the packs to the tankî is the primary reason for

elevated temperatures in those fuel tanks and indicated that one purpose of the proposed

rule changes (to ìminimizeî the flammability of fuel tanks in newly designed aircraft) is

ìto preclude the future use of such design practices.î In the NPRM, the FAA also

recognized that airplanes without heat sources adjacent to fuel tanks have significantly

reduced levels of flammability exposure (5 percent of the fleet operational time) when

650 It should be noted that the potential for 747 CWT overheating and methods for mitigating against it

were studied by Boeing and documented in a 1980 Fuel Tank Heating Study prepared for the USAF in

connection with the E-4B, a military variant of the 747. According to Boeing, its personnel associated with

the TWA flight 800 investigation did not become aware of the report until late in the investigation. The

Safety Board recognizes that the military variant of the 747 is not directly comparable to the civilian 747 and

that the focus of that study was fuel pump functionality, not flammability. Nonetheless, it is unfortunate that

potentially relevant information about 747 CWT overheating and corrective measures was not provided to

the FAA or to 747 operators earlier. According to Boeing, all of its company reports and resources are now

electronically accessible through key-word searches, and its accident investigation processes have been

revised and now include the ability to electronically search the companyís technical documents. 
Analysis 300 Aircraft Accident Report

compared with airplanes that have heat sources adjacent to fuel tanks (30 percent of the

fleet operational time).651

The Safety Board is pleased that the FAA agrees that the heat generated by the

packs located under the CWT significantly increases the flammability exposure in the

CWT. Further, the Safety Board concludes that the placement of heat-generating

equipment under a fuel tank containing Jet A fuel can unnecessarily increase the amount

of time that the airplane is operating with a flammable fuel/air mixture unless measures

are in place to either (1) prevent the heat from entering the CWT or (2) eliminate the

flammable vapors inside the CWT. 

Although the Safety Board supports the FAAís efforts to address fuel tank

flammability in new designs, the Board is concerned about the FAAís apparent premise

that minimizing, rather than eliminating, fuel tank flammability is an acceptable goal. The

ARAC working group (the FTHWG) established to advise the FAA in its rulemaking

efforts concluded that the safety record of fuel tanks located in the wings (which the

FTHWG calculated were flammable about 7 percent of the fleet operational time) was

adequate and that if the same level of safety could be achieved in CWTs, the overall safety

objective could be achieved. Thus, the FTHWG proposed limiting the exposure to

flammable conditions in all fuel tanks to less than 7 percent of the expected fleet

operational time.

FAA staff indicated during meetings with Safety Board staff on February 29, 2000,

that the FAA did not intend to endorse the FTHWGís proposed 7 percent exposure criteria.

However, the FAAís draft guidance652 concerning compliance with the proposed new

requirement to minimize fuel tank flammability reflects exposure criteria that are tied to

the exposure of wing tanks; these criteria are reiterated in the FAAís May 17, 2000, letter

to the Board describing its actions in response to several safety recommendations,

including A-96-174 and -175. The draft guidance states that ìthe intent of the regulation is

to require that the exposure to formation or presence of flammable vapors is equivalent to

that of an unheated wing tank in the transport airplane being evaluated,î which, as

previously mentioned, the FTHWG calculated to be about 7 percent of the fleet

operational time. 

The Safety Board registered two concerns about such exposure criteria in its

comments on NPRM 99-18. First, the Board stated that because the exposure criteria is

based on a fleetwide average, it does not account for increased risks that may exist at

specific locations, during certain time periods, or for certain flights.653 Second, the Board

651 The FAA amended this estimate (of how much of the fleet operational time CWTs are flammable) to

35 percent in its May 17, 2000, letter to the Safety Board, based on its analysis of recent flammability

research.

652 FAA Draft AC 25.981-2X, ìFuel Tank Flammability Minimization,î dated January 12, 2000.

653 The Safety Board notes that of the 16 fuel tank explosions documented by the FTHWG, 11 occurred

in the months of May, June, and July, and only 2 occurred outside of warm climates or summer months.

Further, short flights in warm climates can have a dramatically higher exposure time than the fleetwide

average. For these flights, attempting to reduce the fleetwide flammability of CWTs to the level of wing

tanks would be of little benefit.
Analysis 301 Aircraft Accident Report

questioned the premise that transport-category airplane wing fuel tanks have an acceptable

safety record, noting that wing fuel tanks have exploded. The Board also stated that the

goal should be to completely eliminate the development of flammable vapors in fuel tanks

to the greatest extent technically feasible (such as would result from the use of inerting

systems). 

In addition, the Safety Board is even more concerned that NPRM 99-18 did not

propose any regulatory changes that address fuel tank flammability in current designs and

in the existing fleet. This is especially disturbing because some operational measures that

can reduce current levels of flammable vapors (such as limiting the on-ground operating

time of air conditioning packs and instead using a ground-based cool air supply and

cooling or ventilating the pack bay) could be accomplished immediately.

The FAA indicated in its May 17, 2000, letter to the Safety Board that it is

evaluating methods for reducing flammability in both future and existing designs,

including directed ventilation and on-board or ground-based fuel tank nitrogen inerting

systems. However, according to the FAA, fuel tank temperature data provided (to date) by

airplane manufacturers in response to its request indicate that directed ventilation systems

would not reduce the temperature of heated CWTs as much as predicted by the FTHWG.

The FAA indicated that it would not reach a final decision regarding the benefits of this

design change until it had received and evaluated all of the requested fuel tank

temperature data.

With regard to fuel tank inerting, the FAA letter stated that its research on

ground-based inerting indicated that the costs associated with the needed ground

equipment are significantly less than what was estimated by the FTHWG and, further, that

analysis indicated that ground-based inerting would reduce the fleetwide average

flammability exposure for affected CWTs to about 2 percent. The FAA indicated that it

planned to conduct ground and flight tests to determine how long ground-based inerting

would remain effective during flight and whether vent system modifications would be

required to ensure its effectiveness. The FAA stated that it is also establishing an ARAC

working group to gather additional data on the practicality of both ground-based and

on-board inerting methods. 

In addition, the FAA indicated that it is evaluating the use of ground sources for

conditioned air (instead of using the air conditioning packs) as an additional measure to

reduce the flammability of CWTs in existing designs and the current fleet. According to

the FAA, use of ground sources on days when the temperature exceeds 60° F would

reduce the fleetwide flammability of CWTs from about 35 percent to about 25 percent.

The FAA stated that a Boeing service letter (SL) recommends that operators use ground

sources for conditioned air when available and practical and that it intends to encourage

operators to follow this recommendation. The Safety Board is pleased by the significant

reduction in flammability that can be achieved by using ground sources for conditioned air

and notes that a requirement that such sources be used seems to be a logical step toward

satisfying the intent of Safety Recommendation A-96-175. 
Analysis 302 Aircraft Accident Report

Although the use of ground sources for conditioned air can provide meaningful

short-term benefits, fuel tank inerting appears to be a more promising, near-term method

that could even more dramatically reduce fuel tank flammability in the existing fleet.654

Thus, the Safety Board is pleased that the FAA has recently begun to aggressively study

this method. However, given that 3 1/2 years have passed since the Board recommended

that the FAA give significant consideration to the issue of fuel tank inerting as a means of

precluding the operation of transport-category airplanes with flammable vapors in the fuel

tanks, and given the nature of the results that have been achieved by prior ARACs

(particularly the FTHWG)655 the Board is concerned that the FAA has chosen to address

this issue by forming another ARAC working group to review and advise the FAA

regarding the practicality of fuel tank inerting methods. The Board strongly encourages

the FAA to consider a broader range of inerting technologies656 and expedite the pace of

its research and rulemaking initiatives with regard to fuel tank inerting. However, based

on the FAAís recent input to the Board, and pending additional expedited action with

regard to fuel tank inerting, Safety Recommendations A-96-174 and -175 are classified

ìOpenóAcceptable Response.î 

2.5 Maintenance and Aging Issues

As previously discussed, no evidence was found that preexisting airplane fatigue

or corrosion contributed to the accident.657 Although small fatigue cracks were found in

the lower chord of the front spar and in shear ties for the floor beams and stiffeners at the

front spar, they would not have degraded the capability of the front spar to react to wing

bending loads as a part of the WCS, nor would they have affected the pattern or severity of

the airplaneís structural breakup. These fatigue cracks appear to have been exposed by

fractures generated by the forces associated with the overpressure event within the CWT

and by the continued breakup of the airplane. No fatigue cracking was associated with the

initial failure of the CWT at the upper end of SWB3. 

However, several potentially unsafe conditions were found in and near the

electrical wiring from the accident airplane, including cracked wire insulation, metal

654 The Safety Board recognizes that fuel tank flammability can be counteracted in a number of different

ways. However, inerting appears to be the most direct and effective method because it does not require

monitoring or predicting temperatures inside the fuel tank. 

655 The FTHWG was tasked with recommending regulatory changes that would ìeliminate or

significantly reduce the hazards associated with explosive vapors in transport category airplane fuel tanks

[of]Önew type designs, in-production airplanes and the existing fleet.î 63 Federal Register 3614

(January 23, 1998). However, its final report recommended regulatory changes only with regard to

flammability reduction in new aircraft designs and also recommended further study of methods for

flammability reduction in the existing fleet and current designs. Further, as previously discussed, the Safety

Board disagreed with the FTHWG conclusion that the safety record of wing fuel tanks was adequate and that

the FAA should strive for a similar safety record for fuel tanks adjacent to heat sources. 

656 For example, the FAA should research the effectiveness of inerting gases other than nitrogen and

various methods of gas storage, delivery, and availability.

657 For more information about preexisting fatigue and corrosion in the accident airplane, see section

2.2.1.1.
Analysis 303 Aircraft Accident Report

shavings adhered to a floor beam along which FQIS wires would have been routed,658

other debris, and sulfide deposits. In addition, as previously mentioned, evidence of

several repairs that did not comply with the guidelines in Boeingís SWPM were found on

the accident airplane. Noncompliant repairs included the use of an oversized strain relief

clamp on the terminal block of the No. 1 fuel tank compensator, which did not adequately

secure the wires; numerous open-ended (rather than sealed) wire splices, which exposed

conductors to possible water contamination; several wire bundles containing numerous

wire splices on adjacent wires at the same location; and excessive solder on the connector

pins inside the fuel totalizer gauge, which had connected the pins/wires from the right

wing main fuel tank and the CWT FQIS.

Although some of these conditions may suggest the need for improved

maintenance, the Safety Board found that deterioration, damage, and contamination of

aircraft wiring and related components, and noncompliant repairs, such as those found on

the accident airplane, were common in the airline transport-category airplanes of

numerous operators (especially in the older airplanes) inspected in connection with this

accident investigation.659 Therefore, the Safety Board concludes that the condition of the

wiring system in the accident airplane was not atypical for an airplane of its age and that

the airplane was maintained in accordance with prevailing industry practices. 

Nonetheless, the Safety Board was concerned about the damage and contamination

found on electrical wiring and components during its examinations of transport-category

airplanes, including the accident airplane. These conditions were especially disturbing

because it was apparent from those examinations that a large portion of aircraft wiring is

difficult, if not impossible, to inspect and/or test because of its inaccessibility (as a result

of being confined in wire bundles or blocked by other obstructions). Moreover, the general

nature of current visual wiring inspection criteria is such that wire damage or other

potentially unsafe conditions may not be detected even on visible and accessible portions

of aircraft wiring. The Safety Board concludes that, until recently, insufficient attention

has been paid to the condition of aircraft electrical wiring, resulting in potential safety

hazards. 

However, this accident investigation has resulted in a heightened awareness of the

importance of maintaining the integrity of aircraft wiring, and the FAA has now

recognized that current maintenance practices may not adequately or proactively address

aging nonstructural systems. In its July 1998 Aging Transport Non-Structural Systems

Plan (which was developed after the FAA received the recommendation of the White

House Commission on Aviation Safety and Security and had participated in Safety Board

airplane inspections), the FAA described the results of its evaluation of five

658 The presence of metal shavings is consistent with maintenance records that described compressed air

being used to blow metal shavings off avionics units. (For more information about the accident airplaneís

maintenance records, see section 1.6.3.2.)

659 The widespread existence of such conditions was also corroborated by Boeingís SL 747-SL-20-048

(January 25, 1995), which detailed similar conditions found in numerous 747s. In that SL, Boeing also stated

that it ìbelieve[d] that the wiring on high time 747 airplanes is holding up exceptionally wellî and noted that

ì[w]iring damage is hard to predict but some areas of wiring experience degradation more frequently.î
Analysis 304 Aircraft Accident Report

transport-category airplanes deemed representative of the ìagingî fleet of

transport-category airplanes. The FAA found conditions similar to those found by the

Board during airplane inspections in connection with this investigation, including

deterioration of wiring and related components; stiff and easily cracked wire;

contamination of wire bundles with metal shavings, dust, and fluids; cracked insulation;

corrosion on connector pins; and improper wire installation and repairs. The FAA also

found, as had Board investigators, that wires contained in wire bundles could be difficult

to inspect.

The FAA concluded that current maintenance practices do not adequately address

wiring components, wire inspection criteria are too general, unacceptable conditions are

not described in sufficient detail, repair instructions and data are difficult to extract from

SWPMs, wire replacement criteria may not be adequate, and current incident/maintenance

reporting procedures do not allow for easy identification of failures. The Safety Board

concurs with these conclusions. The FAAís plan also detailed several tasks and associated

subtasks aimed at correcting these problems, including improving wiring inspection

criteria and providing more detailed descriptions of undesirable conditions; improving

inspector training to ensure that it adequately addresses the recognition and repair of aging

wiring components; and developing methods for nondestructive testing of wiring.660

To assist in addressing these problems, the FAA established the Aging Transport

Systems Rulemaking Advisory Committee (ATSRAC), the activities of which the Safety

Board has been following with great interest. The Board is encouraged that other

Government agencies, including the U.S.Navy and USAF, have increased their research

into aging aircraft systems and shared the resulting information and reports with the FAA

and that the White House has formed a Wire Safety Research interagency working group.

Although the FAAís research is scheduled to continue into future fiscal years, the

ATSRAC charter will expire in January 2001, unless it is extended. In light of the short

time remaining for the completion of the ATSRACís work, the Board is concerned that the

ATSRACís final report and recommendations may not fully and adequately address all of

the issues identified in the Aging Transport Non-Structural Systems Plan.

The Safety Board supports the FAAís ongoing efforts to address the

maintenance/aging issues identified by the Aging Transport Non-Structural Systems Plan

and will continue to follow the activities of the ATSRAC and other Government and

industry initiatives related to aircraft maintenance and aging problems. The Safety Board

concludes that the issues defined in the FAAís Aging Transport Non-Structural Systems

Plan are important safety issues that must be fully addressed through appropriate changes,

including rulemaking. Therefore, the Safety Board believes that the FAA should,

regardless of the scope of the ATSRACís eventual recommendations, address (through

rulemaking or other means) all of the issues identified in the Aging Transport

Non-Structural Systems Plan, including the need for improved training of maintenance

660 The Safety Board has participated in demonstration tests of automated test equipment in which faults

were detected that could not be visually seen. Lectromec documented similar results during inspections of

five U.S. Navy P-3 airplanes, finding that visual inspections revealed only 25 to 39 percent of the defects

found by electrical inspections. For more information about these inspections, see section 1.18.2.5.
Analysis 305 Aircraft Accident Report

personnel to ensure adequate recognition and repair of potentially unsafe wiring

conditions; the need for improved documentation and reporting of potentially unsafe

electrical wiring conditions; and the need to incorporate the use of new technology, such

as AFCBs and automated wire test equipment. To determine whether adequate progress is

being made in these areas, the Safety Board believes that, within 90 days, the FAA should

brief the Safety Board on the status of its efforts to address all of the issues identified in

the Aging Transport Non-Structural Systems Plan.
306 Aircraft Accident Report

3. Conclusions

3.1 Findings

1. The flight crew was properly certificated and qualified and had received the training

and off-duty time prescribed by Federal regulations. No evidence indicated any

preexisting medical or behavioral conditions that might have adversely affected the

flight crewís performance during the accident flight.

2. The airplane was certificated, equipped, and dispatched in accordance with Federal

regulations and approved TWA procedures.

3. At the time of the accident, there were light winds and scattered clouds in the area, but

there were no significant meteorological conditions that might have disrupted the

flight.

4. The in-flight breakup of TWA flight 800 was not initiated by a preexisting condition

resulting in a structural failure and decompression. 

5. The in-flight breakup of TWA flight 800 was not initiated by a bomb or a missile

strike. 

6. The fuel/air vapor in the ullage of TWA flight 800ís center wing fuel tank was

flammable at the time of the accident. 

7. A fuel/air explosion in the center wing fuel tank of TWA flight 800 would have been

capable of generating sufficient internal pressure to break apart the tank. 

8. The witness observations of a streak of light were not related to a missile, and the

streak of light reported by most of these witnesses was burning fuel from the accident

airplane in crippled flight during some portion of the postexplosion preimpact

breakup sequence. The witnessesí observations of one or more fireballs were of the

airplaneís burning wreckage falling toward the ocean.

9. The TWA flight 800 in-flight breakup was initiated by a fuel/air explosion in the

center wing fuel tank. 

10. Boeingís design practice that permits parts less than 3 inches long in any direction to

be electrically unbonded may not provide adequate protection against potential

ignition hazards created by static electricity generated by lightning or other

high-energy discharges. 
Conclusions 307 Aircraft Accident Report

11. It is very unlikely that the flammable fuel/air vapor in the center wing fuel tank

(CWT) on TWA flight 800 was ignited by a lightning or meteor strike; a missile

fragment; a small explosive charge placed on the CWT; auto ignition or hot surface

ignition, resulting from elevated temperatures produced by sources external to the

CWT; a fire migrating to the CWT from another fuel tank via the vent (stringer)

system; an uncontained engine failure or a turbine burst in the air conditioning packs

beneath the CWT; a malfunctioning CWT jettison/override pump; a malfunctioning

CWT scavenge pump; or static electricity. 

12. Electromagnetic interference from radio frequency sources external to TWA

flight 800 did not produce enough energy to ignite the fuel/air vapor in the center

wing fuel tank.

13. Electromagnetic interference from personal electronic devices played no role in

igniting the fuel/air vapor in TWA flight 800ís center wing fuel tank.

14. It is unlikely that electromagnetic interference from aircraft system wiring played a

role in igniting the fuel/air vapor in TWA flight 800ís center wing fuel tank.

15. Existing standards for wire separation may not provide adequate protection against

damage from short circuits.

16. A short circuit producing excess voltage that was transferred to the center wing tank

(CWT) fuel quantity indication system wiring is the most likely source of ignition

energy for the TWA flight 800 CWT explosion.

17. Silver-sulfide deposits on fuel quantity indication system components inside fuel

tanks pose a risk for ignition of flammable fuel/air vapor. 

18. The ignition energy for the center wing fuel tank (CWT) explosion most likely

entered the CWT through the fuel quantity indication system (FQIS) wiring, and,

although it is possible that the release of ignition energy inside the CWT was

facilitated by the existence of silver-sulfide deposits on an FQIS component, neither

the energy release mechanism nor the location of the ignition inside the CWT could

be determined from the available evidence.

19. Failure modes and effects analyses and fault tree analyses should not be relied upon

as the sole means of demonstrating that an airplaneís fuel tank system is not likely to

experience a catastrophic failure.

20. A fuel tank design and certification philosophy that relies solely on the elimination of

all ignition sources, while accepting the existence of fuel tank flammability, is

fundamentally flawed because experience has demonstrated that all possible ignition

sources cannot be predicted and reliably eliminated. 

21. Operating transport-category airplanes with flammable fuel/air mixtures in fuel tanks

presents an avoidable risk of an explosion.
Conclusions 308 Aircraft Accident Report

22. The placement of heat-generating equipment under a fuel tank containing Jet A fuel

can unnecessarily increase the amount of time that the airplane is operating with a

flammable fuel/air mixture unless measures are in place to either (1) prevent the heat

from entering the center wing fuel tank (CWT) or (2) eliminate the flammable vapor

inside the CWT. 

23. The condition of the wiring system in the accident airplane was not atypical for an

airplane of its age, and the airplane was maintained in accordance with prevailing

accepted industry practices. 

24. Until recently, insufficient attention has been paid to the condition of aircraft

electrical wiring, resulting in potential safety hazards. 

25. The issues defined in the Federal Aviation Administrationís Aging Transport

Non-Structural Systems Plan are important safety issues that must be fully addressed

through appropriate changes, including rulemaking.

3.2 Probable Cause

The National Transportation Safety Board determines that the probable cause of

the TWA flight 800 accident was an explosion of the center wing fuel tank (CWT),

resulting from ignition of the flammable fuel/air mixture in the tank. The source of

ignition energy for the explosion could not be determined with certainty, but, of the

sources evaluated by the investigation, the most likely was a short circuit outside of the

CWT that allowed excessive voltage to enter it through electrical wiring associated with

the fuel quantity indication system. 

Contributing factors to the accident were the design and certification concept that

fuel tank explosions could be prevented solely by precluding all ignition sources and the

design and certification of the Boeing 747 with heat sources located beneath the CWT

with no means to reduce the heat transferred into the CWT or to render the fuel vapor in

the tank nonflammable.
309 Aircraft Accident Report

4. Recommendations

4.1 New Recommendations

As a result of the investigation of the TWA flight 800 accident, the National

Transportation Safety Board makes the following recommendations to the Federal

Aviation Administration (FAA):

Examine manufacturersí design practices with regard to bonding of

components inside fuel tanks and require changes in those practices, as

necessary, to eliminate potential ignition hazards. (A-00-105)

Review the design specifications for aircraft wiring systems of all

U.S.-certified aircraft and (1) identify which systems are critical to safety

and (2) require revisions, as necessary, to ensure that adequate separation is

provided for the wiring related to those critical systems. (A-00-106)

Require the development and implementation of corrective actions to

eliminate the ignition risk posed by silver-sulfide deposits on fuel quantity

indication system components inside fuel tanks. (A-00-107)

Regardless of the scope of the Aging Transport Systems Rulemaking

Advisory Committeeís eventual recommendations, address (through

rulemaking or other means) all of the issues identified in the Aging

Transport Non-Structural Systems Plan, including

ï the need for improved training of maintenance personnel to ensure

adequate recognition and repair of potentially unsafe wiring

conditions;

ï the need for improved documentation and reporting of potentially

unsafe electrical wiring conditions; and

ï the need to incorporate the use of new technology, such as arc-fault

circuit breakers and automated wire test equipment.

To determine whether adequate progress is being made in these areas, the

Safety Board believes that, within 90 days, the Federal Aviation

Administration should brief the Safety Board on the status of its efforts to

address all of the issues identified in the Aging Transport Non-Structural

Systems Plan. (A-00-108)
Recommendations 310 Aircraft Accident Report

4.2 Previously Issued Recommendations Resulting 

From This Accident Investigation

As a result of the TWA flight 800 accident investigation, the Safety Board issued

the following safety recommendations to the FAA on December 13, 1996:

Require the development and implementation of design or operational

changes that will preclude the operation of transport-category airplanes

with explosive fuel/air mixtures in the fuel tanks: 

(a) Significant consideration should be given to the development of

airplane design modifications, such as nitrogen-inerting systems and the

addition of insulation between heat-generating equipment and fuel tanks.

Appropriate modifications should apply to newly certificated airplanes

and, where feasible, to existing airplanes. (A-96-174)

(b) Pending implementation of design modifications, require modifications

in operational procedures to reduce the potential for explosive fuel/air

mixtures in the fuel tanks of transport-category aircraft. In the 747,

consideration should be given to refueling the center wing fuel tank (CWT)

before flight whenever possible from cooler ground fuel tanks, proper

monitoring and management of the CWT fuel temperature, and

maintaining an appropriate minimum fuel quantity in the CWT. (A-96-175) 

Require that the 747 Flight Handbooks of TWA and other operators of 747s

and other aircraft in which fuel tank temperature cannot be determined by

flight crews be immediately revised to reflect the increases in center wing

fuel tank (CWT) fuel temperatures found by flight tests, including

operational procedures to reduce the potential for exceeding CWT

temperature limits. (A-96-176)

Require modification of the center wing fuel tank of 747 airplanes and the

fuel tanks of other airplanes that are located near heat sources to

incorporate temperature probes and cockpit fuel tank temperature displays

to permit determination of fuel tank temperatures. (A-96-177)

As a result this accident investigation, the Safety Board also issued the following

recommendation to the FAA on February 18, 1997:

Develop and implement procedures, including a checklist of safety-related

items, for the handling and placement of explosive training aids by

K-9 explosives detection teams to prevent contamination of aircraft and

airport facilities and to ensure an effective K-9 explosives detection

program. (A-97-11)
Recommendations 311 Aircraft Accident Report

As a result of this accident investigation, the Safety Board also issued the

following recommendations to the FAA on April 7, 1998:

Issue, as soon as possible, an airworthiness directive to require a detailed

inspection of fuel quantity indication system wiring in Boeing 747-100,

-200, and -300 series airplane fuel tanks for damage, and the replacement

or the repair of any wires found to be damaged. Wires on Honeywell Series

1-3 probes and compensators should be removed for examination.

(A-98-34)

Issue an airworthiness directive to require the earliest possible replacement

of the Honeywell Corporation Series 1-3 terminal blocks used on Boeing

747 fuel probes with terminal blocks that do not have knurled surfaces or

sharp edges that may damage fuel quantity indication system wiring.

(A-98-35)

Conduct a survey of fuel quantity indication systems probes and wires in

Boeing 747s equipped with systems other than Honeywell Series 1-3

probes and compensators and in other model airplanes that are used in Title

14 Code of Federal Regulations Part 121 service to determine whether

potential fuel tank ignition sources exist that are similar to those found in

the 747. The survey should include removing wires from fuel probes and

examining the wires for damage. Repair or replacement procedures for any

damaged wires that are found should be developed. (A-98-36)

Require research into copper-sulfide deposits on fuel quantity indication

system parts in fuel tanks to determine the levels of deposits that may be

hazardous, how to inspect and clean the deposits, and when to replace the

components. (A-98-37)

Require in Boeing 747 airplanes, and in other airplanes with fuel quantity

indication system (FQIS) wire installations that are corouted with wires

that may be powered, the physical separation and electrical shielding of

FQIS wires to the maximum extent possible. (A-98-38)

Require, in all applicable transport airplane fuel tanks, surge protection

systems to prevent electrical power surges from entering fuel tanks through

fuel quantity indication system wires. (A-98-39)

For additional information about these recommendations, see section 1.18.3 of this

report.
Recommendations 312 Aircraft Accident Report

4.3 Previously Issued Recommendations Classified in 

This Report

The following previously issued recommendations are classified in this report:661

ï Safety Recommendation A-96-174 (previously classified ìOpenó

Unacceptable Responseî) is classified ìOpenóAcceptable Responseî in

section 2.4.2 of this report.

ï Safety Recommendation A-96-175 (previously classified ìOpenó

Unacceptable Responseî) is classified ìOpenóAcceptable Responseî in

section 2.4.2 of this report.

ï Safety Recommendation A-98-37 (previously classified ìOpenóAcceptable

Responseî) is classified ìClosedóAcceptable ActionóSupersededî in section

2.3.2.2.3.1 of this report.

ï Safety Recommendation A-98-38 (previously classified ìOpenóAcceptable

Responseî) is again classified ìOpenóAcceptable Responseî in section

2.3.2.1.2.2.3 of this report.

ï Safety Recommendation A-98-39 (previously classified ìOpenóUnacceptable

Responseî) is classified ìOpenóAcceptable Responseî in section

2.3.2.1.2.2.4 of this report.

661 For the text of these recommendations, see section 4.2.

BY THE NATIONAL TRANSPORTATION SAFETY BOARD

James E. Hall

Chairman

John A. Hammerschmidt

Member

John J. Goglia

Member

George W. Black, Jr.

Member

Carol J. Carmody

Member

Adopted August 23, 2000
Appendixes on pages 313-416 are contained in a separate document

for download from the NTSB Web site:

http://www.ntsb.gov/publictn/2000/AAR0003.htm

417 Aircraft Accident Report

Glossary of Terms

Air conditioning pack: An assembly of air conditioning system components that

reduces the temperature and pressure of hot bleed air that is then routed to pressurized

areas of the cockpit, cabin, and cargo compartments to provide environmental control

(pressurization, ventilation, and temperature). The bleed air source can be any one or a

combination of the following: engines, auxiliary power unit (APU), or ground air

connections.

Air cycle machine (ACM): An assembly in the air conditioning system that

includes components such as a fan and compressor. When partially cooled air from the

heat exchangers requires additional cooling, the air is routed through the ACM expansion

turbine for maximum cooling.

Alternating current (a.c.): An electric current that periodically changes in

direction and constantly changes in magnitude.

Ampere (amp): The basic unit of measurement of electric current flow. 

Arcing: Arcing is defined by Underwriters Laboratories, Inc., as a luminous

discharge of electricity across an insulating medium. The electrical discharge of an arc

can involve temperatures of several thousand degrees Celsius.

Auto ignition: Spontaneous ignition of a fuel/air vapor when it is sufficiently

elevated in temperature for ignition to occur without direct contact with an ignition source,

such as a spark, arc, or hot surface or filament. Auto ignition is highly dependent upon

many factors, including the size of the heated volume and container and other

environmental conditions that affect fuel vapor. Existing research indicates that the auto

ignition temperature for Jet A fuel/air vapor at sea level is about 460o Fahrenheit (F).

Auto ignition temperature increases as the altitude increases. 

Ballistic coefficient: The weight of an object divided by the product of its drag

coefficient multiplied by its area, used to determine the motion of an object in an

atmospheric environment. 

Bleed air: The hot pressurized air ducted from an airplaneís engines, APU, or a

ground source for use by other airplane systems (such as pressurization and air

conditioning).

Body station (STA): A longitudinal point along an airplaneís fuselage, identified

numerically by its distance in inches from a reference point. In a classic 747, this point is

90 inches forward of the airplaneís nose. 

Bomb: (as used in this report) An explosive device designed to release destructive

material at high velocity upon detonation, as distinguished from a small explosive charge.
Glossary of Terms 418 Aircraft Accident Report

Bonding: Connecting components to maintain them at a common electric

potential.

Boost pump: (as used in this report) A pump mounted in the wing fuel tanks

designed to move fuel from certain wing tanks to the engines.

Capacitance: The property of conductors separated by a dielectric material (for

example, air or fuel) that permits the storage of electricity when potential differences exist

between conductors. 

Center wing fuel tank (CWT): A fuel tank that, in the 747-100 series, is located

in the wing center section (WCS), between the rear spar and spanwise beam (SWB) 3 and

that has a Jet A fuel capacity of 86,363 pounds (12,890 gallons). (See wing center

section.)

Conductivity: A measure of the extent to which a material is capable of

conducting an electric current. (See ohm.)

Connector: (as used in this report) A device that makes an in-line connection(s)

between one or more wires for a continuous electrical path(s) at a location where the wires

are subject to being disconnected and reconnected without mismatching circuits. Typical

military-specification multicontact electrical connectors are assembled from two

subassemblies, the plug and receptacle, which mate to connect wires with pin and socket

contacts.

Contact: (as used in this report) A device within an electrical connector used to

provide the electrical path joining two individual wires.

Coupling: Transferring energy between elements or circuits of an electrical

system.

Current: The movement of electricity (the flow of electrons) through a conductor.

Measured in amps.

Dielectric: A nonconductor of electricity. 

Direct current: An electric current that flows continuously in one direction.

Dry bay: A compartment in the WCS that is not intended to contain fuel. In the

747-100, a dry bay is located between SWB3 and the forward spar.

Electromagnetic environment (EME): The total of all electromagnetic fields and

the associated frequencies, power levels, and polarizations in a given or defined region.

The EME consists of natural and manufactured sources of electromagnetic energy.

Electromagnetic interference: (as used in this report) Electromagnetic energy

from a source either internal or external to an aircraft that imposes greater-than-intended

voltage on an electrical system. 
Glossary of Terms 419 Aircraft Accident Report

Energy: The capacity for doing work. It may exist in potential, kinetic, thermal,

electrical, chemical, nuclear, or other forms and be transformed from one form to another.

Electromagnetic energy is expressed in units of work, such as joules (J) or kilowatt-hours.

Explosion: (as used in this report) The sudden and rapid escape of gases from a

confined space, accompanied by high temperatures, violent shock, and loud noise. 

Extremely improbable failure condition: As defined in Federal Aviation

Administration Advisory Circular (AC) 25.1309-1A, a condition so unlikely that it is not

anticipated to occur during the entire operational life of all airplanes of one type and that

has a probability of occurrence on the order of 1 x 10-9 or less each flight hour, based on a

flight of mean duration for the airplane type. 

Failure modes and effects analysis: A structured and systematic analytical

method for identifying potential failure modes of a component or system and for

evaluating the potential risk(s) that might be posed by various failure modes. 

Fault tree analysis: An analysis designed to examine an end event through

consideration of assumed precipitating events. Each of these precipitating events is in turn

broken down until a level is attained in which no additional precipitating events will occur.

The total of all of the events and the way in which they are tied together creates the fault

tree analysis.

Fireball: (as defined by the TWA flight 800 investigationís Witness Group) One

or more downward-moving ball(s) of fire in the sky. According to the Witness Groupís

definition, it could be characterized (by witnesses) as either stationary or descending;

however, to meet the groupís definition of a fireball, it must not have been reported to

have appeared in the sky after the termination of a streak of light (if such a streak were

reported). It could not have been an ascending object or an object that met the groupís

definition of a streak of light.

Flash point: The minimum temperature at sea level at which a liquid fuel

vaporizes sufficiently to form an ignitable mixture with air (when exposed to an open

flame), as determined by a standardized test procedure. 

Flashing: (as used in this report) A category of electrical activity observed during

short-circuit tests conducted by Lectromechemical Design Company (Lectromec) as a part

of the TWA flight 800 accident investigation, characterized by an arcing discharge seen as

a single flash of light with an accompanying popping sound.

Fuel mass loading: A measure of the amount of fuel relative to the entire volume

of its container.

Fuel quantity indication system (FQIS): In the 747-100, a system that measures

changes in the capacitance of tubular probes located in each fuel tank for the display of

fuel quantity on cockpit gauges and on repeater gauges located at the fueling station in the

left wing. The system is also connected to systems that require fuel quantity information,
Glossary of Terms 420 Aircraft Accident Report

including the gross weight/total fuel weight indicator, airborne integrated data acquisition,

and the volumetric shutoff ([VSO] ground refueling) systems.

Fuel quantity compensator: A component used in the FQIS or VSO system to

compensate for variations in the dielectric constant of fuel, which varies from one type of

fuel to another (and even within the same type of fuel, depending on the ìbatchî and age

of the fuel) to ensure consistent fuel quantity indications. There is at least one

compensator in each fuel tank.

Fuel quantity probe: A component of the FQIS, a set of which is positioned in

each of an airplaneís fuel tanks, used to measure the quantity of fuel in each tank through

a range of levels within the tank. The 747-100 may have 65 or more fuel quantity probes

positioned within the 7 fuel tanks.

Fuel washing: Motion of fuel over a part, which can lead to electrostatic charge

accumulation or removal of contaminants (such as sulfides).

Green zone: One of three debris fields, labeled during the TWA flight 800

accident investigation, from which the accident airplaneís wreckage was recovered. This

zone was located farthest east (farthest from John F. Kennedy International Airport [JFK])

in the wreckage distribution. This zone contained pieces of wreckage from both wings

and most of the aft portions of the fuselage, including the following: both wings; all four

engines; pieces of SWB1, SWB2, mid spar, and rear spar; the aft portion of the keel beam;

the main landing gear; and the tail section. 

Grounding: Connecting electrical circuits to a large common conductor

considered to be at zero electrical potential such as the earth or, in the case of an airplane,

the skin of the airplane, which is at a fixed electrical potential. 

Heat exchangers: (as used in this report) Devices in the 747-100 used to initially

cool the heated bleed air coming from the engines located within the air conditioning pack

bay under the CWT.

High-intensity radiated fields: High-power electromagnetic fields that exist in a

defined environment, particularly in the vicinity of high-power radar sites, broadcast

antennas, and other high-power radio frequency sources. 

Hot surface ignition: A phenomenon in which a very high temperature surface

comes into contact with fuel or fuel vapor that results in ignition. Hot surface ignition is

highly dependent upon many factors, including the geometric aspects of the hot surface

and other environmental conditions that affect fuel and fuel vapor. Existing research

indicates that hot surface ignition temperatures at sea level for fuels similar to Jet A range

from 900° to 1,300° F. The hot surface ignition temperature increases as the altitude

increases.
Glossary of Terms 421 Aircraft Accident Report

Ignition energy: The quantity of heat or electrical energy that must be absorbed

by a fuel/air vapor mixture in a finite volume to generate a propagating flame. Commonly

measured as the energy provided by a small (millimeter size) spark. 

Impeller: A blade on a rotating part in an air compressor or fuel pump.

Improbable failure condition: As defined in AC 25.1309-1A, a condition not

anticipated to occur during the entire operational life of a single random airplane;

however, it may occur occasionally during the entire operational life of all airplanes of one

type. It has the probability of occurrence on the order of 1 x 10-5 or less, but greater than

on the order of 1 x 10-9 each flight hour, based on a flight of mean duration for the airplane

type. 

Jet A fuel: A kerosene fuel used in civilian turbine engine airplanes. Jet A fuel is

composed of a mixture of more than 100 distinct types of hydrocarbon molecules; the

precise composition often varies between refinery and by season. Jet A fuel is specified to

have a minimum flash point of 100° F. Jet A-1 is a similar fuel, but has a slightly lower

freezing point. Although Jet A fuel is available in some other countries, it is used

primarily in the United States. 

Jettison/override pump: In the 747, a fuel pump that has two uses, one of which

is to jettison fuel overboard when this function is selected at the flight engineering station.

The pump is also designed to provide fuel to the engine manifolds at a higher pressure

than the boost pumps, which are located in the wing fuel tanks. Two CWT

jettison/override pumps are mounted on the rear spar of the CWT, and pressure from these

pumps closes the main tank check valves, resulting in the use of fuel from the CWT before

that of fuel from the wing tanks.

Joule (J): A unit of measurement of electrical work or energy; 1 J is the amount of

work done by 1 watt of power in 1 second.

Keel beam: A box-shaped, load-bearing structure located along the airplaneís

centerline that extends from the aft wall of the forward cargo compartment below the

WCS and through the main and body landing gear compartments to the forward wall of

the aft cargo compartment. The CWT is located above the keel beam.

Kilojoule (kJ): A unit of measurement equaling 1,000 J.

Lower flammability limit: The lowest temperature that will provide a sufficient

concentration of fuel vapor to propagate a flame.

Millijoule (mJ): A unit of measurement equaling one-thousandth of a J.

Minimum ignition energy (MIE): The minimum quantity of heat or electrical

energy that must be absorbed by an optimal fuel/air vapor mixture in a finite volume to

generate a propagating flame. The MIE for Jet A fuel is generally accepted to be about

0.25 mJ.
Glossary of Terms 422 Aircraft Accident Report

Ohm: The unit of resistance of an electrical conductor, at which the fall of

potential is 1 volt when the current is 1 amp. (See resistance.)

Overpressure event: (as used in this report) An event in which the pressure in the

CWT is increased in a relatively short time to a level at which the structural integrity of

the CWT is compromised.

Power: The time rate of energy transfer; the practical unit of measurement is

1 watt. (See watt.) 

Primary radar target: A radar target produced when a radar signal reflects off of

an objectís surface and returns to a ground-based radar antenna/site for processing and

display.

Quenching: The extinguishment of a combustion flamefront; often as a result of

decreased temperature or propagation through a passageway, such as an orifice or a vent. 

Raceway: A term used to refer to areas within the 747 where wire bundles are

grouped into a common route.

Red zone: One of three debris fields, labeled during the TWA flight 800 accident

investigation, from which the accident airplaneís wreckage was recovered. This zone was

located farthest west (closest to JFK) in the wreckage distribution. A relatively small

amount of widely dispersed debris was recovered from the red zone, including the

following: pieces from the WCS front spar and SWB3, the manufacturing access door

from SWB2, pieces of the fuselage from STA 840 and STA 1000, main cabin floor beams

and flooring material from above and in front of the WCS area, and the two forward air

conditioning packs. (See figure 3a.) 

Resistance: The property of a conductor that tends to restrict the flow of an

electric current. (See ohm.)

Ring chord: An angle member that attaches the bottom of the forward fuselage

section to the front side of the lower pressure bulkhead and the front spar.

Scavenge pump: (as used in this report) A small fuel pump designed to remove

the last amounts of accessible fuel from the lowest point of a 747 CWT and discharge the

fuel into the left inboard fuel tank. Although the scavenge pump removes fuel not

accessible by the jettison/override pumps, a small amount of residual fuel will remain in

the fuel tank that the scavenge pump is not able to remove.

Scintillation: (as used in this report) A category of electrical activity observed

during short-circuit tests conducted by Lectromec as a part of the TWA flight 800 accident

investigation, characterized by a high frequency micro-discharge that usually results in the

formation of char or soot on a wire (and adjacent wires) over time. 
Glossary of Terms 423 Aircraft Accident Report

Secondary radar target: A radar target produced when a radar signal is detected

by an airplaneís transponder, which transmits a coded message in response to interrogation

by a ground-based transmitter. 

Shielding: (as used in this report) Metal covers placed around electric wires and

electronic devices to prevent the intrusion of external electrostatic and electromagnetic

fields.

Short circuit: An unintended current path between conductors. Short circuits can

occur either directly, if the protective insulation covering between internal conductors in

each wire is compromised and there is direct contact between the conductors, or through a

ìbridgeî created by contaminants, such as metal shavings or fluid.

Sleeve (wire): A woven or flexible jacket that protects electrical wiring. 

Small explosive charge: A minimal amount of a highly explosive material (such

as a plastic explosive) detonated by a fusing device. 

Spanwise beam 1: One of the lateral beams in the CWT that divides it into

compartments and supports the structure of the WCS. SWB1 is located between the mid

and rear spars.

Spanwise beam 2: One of the lateral beams in the CWT that divides it into

compartments and supports the structure of the WCS. SWB2 is located between the mid

spar and SWB3. 

Spanwise beam 3: One of the lateral beams in the CWT that divides it into

compartments and supports the structure of the WCS. SWB3 is located between SWB2

and the front spar.

Spar: A beam that extends laterally through the WCS into the outboard wing

structure. The 747-100 has front, mid, and rear spars.

Splice (wire): (as used in this report) A fixed connection of two electrical wires.

Splices can be made by various methods, such as soldering wires together or with crimped

metal barrels, and are typically covered by insulating material.

Streak of light: (as defined by the TWA flight 800 investigationís Witness Group)

An object moving in the sky that could be variously described in witness documents as a

point of light, fireworks, a flare, a shooting star, or something similar, which was usually

described as ascending, but could also be described as arcing over and/or descending.

Streak-of-light witness: (as defined by the TWA flight 800 investigationís

Witness Group) Any witness who reported seeing an airplane in the general vicinity of a

streak of light at the same time that the streak of light was visible (around the time and

vicinity of the TWA flight 800 accident). (See witness.)
Glossary of Terms 424 Aircraft Accident Report

Stringer: A stiffening member found in the 747-100ís fuselage and wings that

helps to support and reinforce the structure.

Strong arcing: (as used in this report) A category of electrical activity observed

during short-circuit tests conducted by Lectromec as a part of the TWA flight 800 accident

investigation, characterized by an arcing discharge that could continue for hundreds of

cycles, typically involving 5 kJ of electrical energy.

Surge tank protection system: A system to detect and extinguish fire that

consists of a series of optical photocells that trigger the discharge of Halon (a fire

extinguishing agent) into the surge tank when a flame or bright light source is sensed in

the surge tank. Discharge of the extinguishing agent is designed to occur about

1 millisecond after the photocell senses a flame. 

Transient suppression device: An electrical device that limits the amount of

energy or current that can pass through it to a predetermined amount.

Transient voltage: A temporary voltage surge or excursion; for example, that

which occurs when first turning an electrical system on. 

Transponder: The airborne receiver/transmitter portion of a radar system that

responds to interrogation signals received from ground-based equipment.

Ullage: The space in a fuel containing tank not occupied by liquid fuel.

Upper flammability limit: The maximum temperature at which a concentration

of fuel vapor will propagate a flame.

Volt: The basic unit of measurement of electromotive force (the force that causes

electrons to flow through a conductor). One volt is the electromotive force required to

cause current to flow at the rate of 1 amp through a resistance of 1 ohm. One volt equals

1 amp times 1 ohm.

Watt: The basic unit of measurement of power. In electrical application, 1 watt

equals 1 volt times 1 amp. (See power.)

Weathering: The change in a liquid fuel chemical composition as a result of

exposure to environmental conditions. An example involves heating and pressure changes

to a vented aircraft fuel tank, where preferential evaporation of the lower molecular

weight components of the jet fuel occurs, resulting in a redistribution of the chemical

composition of the remaining liquid fuel.

Wing center section: A large structural box located aft of the forward cargo

compartment and forward of the main landing gear bay in the lower fuselage between the

wings, which comprises the CWT and a dry bay directly forward of the CWT. (See center

wing fuel tank.)
Glossary of Terms 425 Aircraft Accident Report

Witness: (as defined by the TWA flight 800 investigationís Witness Group)

Anyone who reported hearing a sound and/or seeing an event or object or objects

(including smoke or fire) in the sky around the time and vicinity of the TWA flight 800

accident. According to the Witness Groupís definition, it must have been likely that the

sound or object observed was related to the crash, and the report must not have been a

secondhand account.

Yellow zone: One of three debris fields, labeled during the TWA flight 800

accident investigation, from which the accident airplaneís wreckage was recovered. This

zone was the smallest of the three zones and was contained within the red zone on its

northeastern side and located in a small concentrated area. This zone contained wreckage

from the forward portion of the fuselage, from STA 90 (the nose of the airplane) to

STA 840, including the cockpit, section 41, and the forward portions of section 42. (See

figure 3a.) This wreckage was found relatively intact.