DOT/FAA/AR-04/41 

Office of Aviation Research 

Washington, D.C. 20591 

Evaluation of Fuel Tank 

Flammability and the FAA Inerting 

System on the NASA 747 SCA 

Michael Burns 

William M. Cavage 

Robert Morrison 

Steven Summer 

December 2004 

Final Report 

This document is available to the U.S. public 

Through the National Technical Information 

Service (NTIS), Springfield, Virginia 22161. 

U.S. Department of Transportation 

Federal Aviation Administration
NOTICE 

This document is disseminated under the sponsorship of the U.S.

Department of Transportation in the interest of information exchange. The 

United States Government assumes no liability for the contents or use 

thereof. The United States Government does not endorse products or 

manufacturers. Trade or manufacturer's names appear herein solely

because they are considered essential to the objective of this report. This 

document does not constitute FAA certification policy. Consult your local 

FAA aircraft certification office as to its use. 

This report is available at the Federal Aviation Administration William J.

Hughes Technical Center's Full-Text Technical Reports page:

actlibrary.tc.faa.gov in Adobe Acrobat portable document format (PDF). 
Technical Report Documentation Page

1. Report No.

DOT/FAA/AR-04/41

2. Government Accession No. 3. Recipient's Catalog No.

 4. Title and Subtitle 

EVALUATION OF FUEL TANK FLAMMABILITY AND THE FAA INERTING 

SYSTEM ON THE NASA 747 SCA

5. Report Date

December 2004 

6. Performing Organization Code

7. Author(s) 

Michael Burns, William M. Cavage, Robert Morrison, and Steven Summer

8. Performing Organization Report No. 

DOT/FAA/AR-04/41 

9. Performing Organization Name and Address 10. Work Unit No. (TRAIS) 

Federal Aviation Administration 

William J. Hughes Technical Center 

Airport and Aircraft Safety 

Research and Development Division 

Fire Safety Branch 

Atlantic City International Airport, NJ 08405

11. Contract or Grant No.

13. Type of Report and Period Covered

Final Report 

12. Sponsoring Agency Name and Address 

U.S. Department of Transportation 

Federal Aviation Administration

Office of Aviation Research 

Washington, DC 20591

14. Sponsoring Agency Code

ANM-112 

15. Supplementary Notes

16. Abstract

Extensive development and analysis has illustrated that fuel tank inerting could, potentially, be cost-effective if air separation 

modules, based on hollow-fiber membrane technology, could be packaged and used in an efficient way. To illustrate this, the 

Federal Aviation Administration (FAA) has developed a prototype onboard inert gas generation system that uses aircraft bleed air 

to generate nitrogen-enriched air (NEA) at varying flows and purities during a commercial airplane flight cycle. A series of 

ground and flight tests were performed, in conjunction with National Aeronautics and Space Administration (NASA) aircraft 

operations personnel, designed to evaluate the FAA inerting system used in conjunction with a compartmentalized center wing 

tank (CWT). Additionally, the flammability of both the CWT and one inboard wing fuel tank was measured. The system was 

mounted on a Boeing 747, operated by NASA, and used to inert the aircraft CWT during testing. The inerting system, CWT, and 

the number 2 main wing tanks were instrumented to analyze the system performance, fuel tank inerting, and flammability. 

The results of the testing indicated that the FAA prototype inerting system operated as expected. Using a variable-flow 

methodology allowed a greater amount of NEA to be generated on descent when compared to the simple dual-flow methodology, 

but it had no measurable effect on the resulting average ullage oxygen concentration after each test, while improving inert gas

distribution by decreasing the worst bay oxygen concentration when three similar tests were compared. The highest average 

ullage oxygen concentration observed on any flight test correlates directly with the worst bay oxygen concentration, illustrating 

the importance of maintaining a low average ullage oxygen concentration in good inert gas distribution. Oxygen diffusion 

between the bays of the tank was relatively rapid, and overnight dispersion of the ullage oxygen concentration was measured to be 

very small. Flammability measurements showed trends very similar to what was expected based on both experimental and 

computer model data. The equilibrium data agreed favorably with data from both the Fuel Air Ratio Calculator and the 

Condensation Model, while transient data trends matched closely with the Condensation Model with some discrepancies in total 

hydrocarbon concentration magnitude at altitude. 

17. Key Words 

Nitrogen-enriched air, OBIGGS, Hollow-fiber membrane, Air 

separation module, Total hydrocarbon concentration, Oxygen 

concentration, Thermocouple, Pressure transducer

18. Distribution Statement 

This Document is available to the public through the National 

Technical Information Service (NTIS), Springfield, Virginia 

22161

19. Security Classif. (of this report)

 Unclassified 

20. Security Classif. (of this page)

 Unclassified

21. No. of Pages

 75

22. Price

Form DOT F 1700.7 (8-72) Reproduction of completed page authorized
ACKNOWLEDGMENT

The Federal Aviation Administration (FAA) and the Airport and Aircraft Safety Research and

Development Division, Fire Safety Branch would like to acknowledge the contributions of

NASA Johnson Space Center in supporting the extensive installation of equipment and 

instrumentation to accomplish the very difficult task of studying in-flight flammability and fuel 

tank inerting during flight tests on the NASA 747 SCA. Additionally, acknowledgement goes to

the excellent test pilots and crew that supported the lengthy and time-consuming flight test plan 

provided by the FAA. The dedication and professionalism of the entire test and support 

personnel at Ellington Field, particularly the zero-g team, which allowed the FAA to achieve the 

project goals and objectives was significant, and their technical know-how and professionalism 

are the primary reasons for the success of the test program. 

Further acknowledgement is extended to NASA Glen Research Center Fire Safety Research 

Program for lending financial and moral support to the project when it was needed most. 

iii/iv
TABLE OF CONTENTS 

Page 

EXECUTIVE SUMMARY xi 

1. INTRODUCTION 1 

1.1 Background 1 

1.2 Scope 2 

2. TEST ARTICLE AND PROCEDURES 2 

2.1 Test Article 2 

2.1.1 Aircraft 2 

2.1.2 Instrumentation 5 

2.1.3 Aircraft Cabin Installation 12 

2.2 Test Procedures 16 

3. ONBOARD INERT GAS GENERATION SYSTEM 18 

3.1 System Flow 18 

3.2 System Interfaces 20 

3.2.1 System Mounting 20 

3.2.2 Mechanical Interface 20 

3.2.3 Electrical Interface 22 

3.3 System Operation 22 

4. ANALYSIS 23 

4.1 Inerting Calculations 24 

4.1.1 Bleed Air Consumption 24 

4.1.2 Average Fuel Tank Oxygen Concentration 24 

4.1.3 Inerting Model 24 

4.2 Flammability Calculations 25 

4.2.1 Estimation of the Vapor Generation in the CWT 25 

4.2.2 Fuel Air Ratio Calculator 25 

5. DISCUSSION OF RESULTS 26 

v
5.1 Fuel Tank Inerting 26 

5.1.1 Inerting System Performance 26 

5.1.2 Ullage Oxygen Concentration Data 34 

5.2 Fuel Tank Flammability 42 

5.2.1 General Flammability Trends 42 

5.2.2 Temperature Effects 43 

5.2.3 Cross-Venting Effects 45 

5.2.4 Comparison of Calculations 48 

6. SUMMARY OF FINDINGS 52 

7. REFERENCES 52 

APPENDICES 

A—Center Wing Tank Instrumentation Panel Installation Drawing

B—Heated Gas Sample Fittings Installation Drawing 

C—Instrumentation Diagram With Channel Listing

D—Surface-Mounted Thermocouple Drawing

E—Onboard Inert Gas Generation System Mechanical Drawing With Interface

Illustration

F—Onboard Inert Gas Generation System Inert/Divert Tubing Assembly Top Drawing

 G—Nitrogen-Enriched Air Deposit Nozzle Installation Drawing

H—Onboard Inert Gas Generation System Wiring Diagram

I—Test Fuel Flashpoint and Distillation Data

vi
LIST OF FIGURES

Figure Page 

1 NASA 747 SCA Aircraft With Space Shuttle Orbiter 3 

2 NASA 747 SCA Fuel Tank Diagram 3 

3 Modified Rear Spar Purge Door 4 

4 (a) Rack-Mounted FAS Installation and (b) NDIR Analyzer 7 

5 The FAS Sample Stream Flow Schematic 7 

6 Top View of a 747-100 CWT With Sample Port Locations 8 

7 Gas Sample Float Valve Installation 9 

8 Top View of a 747-100 CWT With Thermocouple Locations 10 

9 Top View of a 747-100 Wing Tank With Thermocouple Locations 11 

10 Master DAS Display 13 

11 Power Distribution for the NASA 747 SCA Cabin 14 

12 Modified Waste Shoot Used as a Conduit at the Aft Pack Bay Area Bulkhead 15 

13 Diagram of Instrumentation Installation in the NASA 747 SCA Cabin 16 

14 System Block Diagram 18 

15 Three-Dimensional Rendering of the OBIGGS Installed in the Pack Bay 20 

16 Onboard Inert Gas Generation System Installed in Test Aircraft Pack Bay 22 

17 Front Panel of OBIGGS Control Box 22 

18 System Performance Data for a Typical Flight Test 26 

19 Comparison of ASM Pressure During the Inception of Three Different Flight 

Tests With Altitude 27 

20 Comparison of NEA Purity During the Inception of Three Different Flight Tests 

With Altitude 28 

21 Comparison of NEA Flow During the Inception of Three Different Flight Tests 

With Altitude 28 

vii
22 Comparison of ASM Input and Output Conditions During the Inception of Three 

Different Flight Tests 29 

23 Comparison of ASM Pressure During the Descent Phase of Three Different Flight

Tests With Altitude 30 

24 Comparison of NEA Purity During the Descent Phase of Three Different Flight 

Tests With Altitude 30 

25 Comparison of NEA Flow During the Descent of Three Different Flight Tests 

With Altitude 31 

26 Comparison of Bleed Air Consumption During the Inception of Three Different 

Flight Tests With Altitude 32 

27 Comparison of Bleed Air Flow During the Descent of Three Different Flight 

Tests With Altitude 32 

28 Correlation of NEA Flow With ASM Pressure for the System Low-Flow Mode 

for Three Different Flight Tests 33 

29 Correlation of NEA Flow With ASM Pressure for the System Variable/High-Flow 

Mode for Three Different Flight Tests 34 

30 Oxygen Concentration Data From all Six Bays With Test Altitude for a Typical 

Flight Test 35 

31 Average Ullage Oxygen Concentration Data Calculated From Six Bay 

Measurements With Test Altitude for a Typical Flight Test 35 

32 Comparsion of Average Ullage Oxygen Concentration With Four Different Test 

Descent Profiles 36 

33 Comparsion of the Worst Bay Oxygen Concentration With Four Different Test 

Descent Profiles 37 

34 Correlation of the Peak Worst Bay Oxygen Concentration With the Peak Average 

Ullage Oxygen Concentration for Four Different Test Descent Profiles 37 

35 Average Ullage Oxygen Concentration Data for a Typical Flight Test Compared 

to a Simple Ullage Inerting Model 38 

36 Oxygen Concentration Data for the Six Bays of the CWT During the Descent of a 

Typical Flight Test 39 

37 Variation in the Oxygen Concentrations in the Six Bays of the CWT During 

Three Flight Tests 40 

viii
38 Comparsion of the Ullage Oxygen Concentration for all Six Bays During a 

3-Hour Turnaround 41 

39 Comparsion of Oxygen Concentration for all Six Bays During a 24-Hour 

Overnight Period With a Fuel Transfer 42 

40 Flight Test 5 THC and Ambient Pressure Measurements 43 

41 Flight Test 5 CWT THC and Temperature Measurements 44 

42 Flight Test 5 Wing Tank THC and Temperature Measurements 44 

43 Flight Test 4 CWT THC and Temperature Measurements 45 

44 Flight Test 0 CWT THC and Ambient Pressure Measurements 46 

45 Flight Test 6 CWT THC and Ambient Pressure Measurements 46 

46 Flight Test 0 CWT Temperature Measurements 47 

47 Flight Test 6 CWT Temperature Measurements 47 

48 Measured and Computed Equilibrium THC Values 49 

49 Center Wing Tank THC Measured and Computed Results for a Ground Test 50 

50 Measured and Calculated THC Measurements for Flight Test 3 50 

51 Measured and Calculated THC Measurements for Flight Test 4 51 

52 Measured and Calculated THC Measurements for Flight Test 5 51 

LIST OF TABLES 

Table Page 

1 Wing Tank Thermocouple Locations 11 

2 Table of Inerting Flight Tests on the NASA 747 SCA 17 

ix
LIST OF ACRONYMS

ACM Air cycle machines 

ASM Air separation modules 

CWT Center wing tank 

DAS Data acquisition system

FAA Federal Aviation Administration 

FAR Fuel air ratio 

FAS Flammability analysis system

ID Inner diameter 

LOC Limiting oxygen concentration 

NASA National Aeronautics and Space Administration

NEA Nitrogen-enriched air 

OBIGGS Onboard inert gas generation system

OBOAS Onboard oxygen analysis system

OEA Oxygen-enriched air 

SCA Space Shuttle Orbiter Carrier Aircraft

SCFM Standard cubic feet per minute 

SOV System shutoff valve 

STA Station 

THC Total hydrocarbon concentration 

WS Wing station 

x
EXECUTIVE SUMMARY

Significant emphasis has been placed on fuel tank safety since the TWA Flight 800 accident in 

July 1996. This has prompted the Federal Aviation Administration (FAA) to study methods that 

could limit the flammability exposure of the commercial transport fleet. The effort was focused

on high-flammability exposure fuel tanks, which are center wing and body-style fuel tanks. 

Extensive development and analysis has illustrated that fuel tank inerting during aircraft 

operation could, potentially, be cost-effective if air separation modules (ASM) could be

integrated into a system and used in an efficient manner. These ASMs are made of thousands of 

tiny, hair-sized, hollow-fiber membranes, which are fabricated into a vessel. When supplied 

with pressurized air, these modules will ventilate a waste stream of gas from a permeate vent that 

is rich in oxygen, carbon dioxide, and water vapor. This allows the product gas passing through 

the ASM to be rich in nitrogen. 

To demonstrate using hollow-fiber membrane ASMs for inerting commercial transport airplane

fuel tanks, the FAA, with the assistance of several aviation-oriented companies, has developed a 

prototype onboard inert gas generation system with ASMs that uses aircraft bleed air to generate 

nitrogen-enriched air (NEA) at varying flows and purities (NEA oxygen concentration) during a

commercial airplane flight cycle. This system was designed to maintain an oxygen concentration 

below 12% in a Boeing 747 center wing tank (CWT) during typical commercial transport 

airplane operations. A series of ground and flight tests were performed, in conjunction with 

National Aeronautics and Space Administration (NASA) aircraft operations personnel, designed

to evaluate the simplified inerting system and examine inerting of a compartmentalized fuel tank. 

Additionally, the flammability of both the center wing and one inboard wing fuel tank was

examined. The FAA inerting system was mounted in the pack bay of a modified 747, operated 

by NASA for Space Shuttle Orbiter transportation, and used to inert the aircraft CWT during

testing. The inerting system, CWT, and the number 2 wing tanks were instrumented to analyze 

system performance, inerting capability, and fuel tank flammability. 

The results of the testing indicated that the FAA inerting system operated as expected. Any 

deviations in system performance from test to test could be explained by the difference in system 

warmup times on the day of the test. Using a variable flow methodology allowed for a greater 

amount of NEA to be generated on descent at a higher oxygen concentration (lower purity) as 

intended, but it had no measurable effect on the resulting average ullage oxygen concentration

after each test when compared to a simple dual-flow system methodology. It did seem to 

improve inert gas distribution by decreasing the worst bay oxygen concentration. The highest 

average ullage oxygen concentration observed on any flight test correlated directly with the 

worst bay oxygen concentration. Oxygen diffusion between the bays of the tank was relatively 

rapid, and the overnight increase of the ullage oxygen concentration was measured to be very 

xi
small. Flammability measurements from both the CWT and the wing tank showed trends very

similar to what was expected based on both experimental and computer model data. The 

equilibrium data agreed favorably with data from both the Fuel Air Ratio Calculator and the 

Condensation Model, while transient data trends matched closely with the Condensation Model, 

with some disagreement with flammability magnitudes at altitude. The measurements generated 

in these flight tests have been used to enhance the capability of these existing flammability 

models and will be used in the future to further improve the predictive flammability calculations.

xii
1

1. INTRODUCTION.

1.1 BACKGROUND. 

Significant emphasis has been placed on fuel tank safety since the TWA Flight 800 accident in 

July 1996. This has prompted the Federal Aviation Administration (FAA) to study methods that 

could limit the flammability exposure of the commercial transport fleet. The effort was focused 

on center wing and body-style fuel tanks that have been identified as being potentially hotter 

during ground operations and more flammable in general [1]. Extensive development and 

analysis has illustrated that flammability reduction using fuel tank inerting during aircraft 

operation could, potentially, be cost-effective if air separation modules (ASM) could be 

integrated into a system and used in an efficient manner. To illustrate this, the FAA, with the 

assistance of several aviation-oriented companies, developed a prototype onboard inert gas 

generation system (OBIGGS) with ASMs that used aircraft bleed air to generate nitrogenenriched air (NEA) at varying flows and purities (NEA oxygen concentration) during a 

commercial airplane flight cycle. This system was designed to maintain an oxygen concentration 

below 12% in a Boeing 747 center wing tank (CWT) during typical commercial transport 

airplane operations. An ASM is made up of thousands of tiny, hair-sized, hollow-fiber 

membranes, which are fabricated into a vessel. When supplied with pressurized air, these 

modules will ventilate a waste stream of gas from a permeate vent that is rich in oxygen, carbon 

dioxide, and water vapor [2]. This allows the product gas passing through the ASM to be rich in 

nitrogen. 

Early ASM OBIGGS work performed by the Department of Defense culminated in a study that 

highlighted the favorable life cycle costs of an on-demand (inert gas generation) system over a 

stored inerting agent system and explosive suppressant foam. This study concluded that an 

inerting system using ASMs with hollow-fiber membrane technology was cost-effective 

compared to other OBIGGS technologies with a relative performance increase of a factor of ten 

over previous ASM technology systems [3]. 

The initial FAA fuel tank inerting flight tests were performed in conjunction with The Boeing 

Company to examine the ability of a ground-based supply of nitrogen (ground-based inerting) to 

inert a commercial airplane CWT. This allowed for validation of the FAA’s Onboard Oxygen 

Analysis System (OBOAS) during flight-testing, which was developed for the project and could 

continuously measure the oxygen concentration of a commercial transport fuel tank ullage using 

conventional laboratory oxygen analyzers and a pressure-controlled sample train [4]. Recent 

FAA flight tests were performed in conjunction with Airbus to test the FAA inerting system, 

which was mounted in the cargo bay of an A320 operated by Airbus, for the purpose of research 

and development. The system performance was degraded to examine the ability of a single ASM 

and two ASMs to inert the CWT of the test aircraft and to help develop a system performance 

model. The basic dual-flow system concept was validated and system scaling was examined. 

The effects of fuel and inert gas distribution were also studied, and a simply ullage inerting 

model was developed to calculate ullage oxygen concentration given a system performance and 

flight cycle [5]. 

Much of the flammability research performed by the FAA has focused on the determination of 

the limiting oxygen concentration (LOC), the oxygen content below which ignition of jet fuel 
2

vapors can no longer be supported. The LOC is the main design criteria for any inerting system, 

because it determines the oxygen levels required to provide adequate fuel tank explosion 

protection. Traditionally, the military has used a value of 9 percent oxygen by volume as the 

inert limit of a fuel tank ullage. An experimental investigation performed by the FAA using a 9-

cubic-foot test article illustrated, however, that no measurable pressure rise was observed when a 

spark was used to ignite a flammable ullage with an oxygen concentration less than 12 percent 

by volume at sea level and 10,000 feet. The LOC value increased in a linear fashion to 

approximately 14.5 percent oxygen at 40,000 feet [6]. The experiment used a small fuel pan, 

radiantly heated from underneath, with the test article in a vacuum chamber to study the ability 

of a simulated ullage to react to a spark source applied under a variety of conditions. Follow-on 

experiments studied a wide variety of fuel/air mass ratios and ignition sources/energies with no 

change in the resulting LOC. A review of previously published literature showed very good 

agreement with measured results. 

1.2 SCOPE. 

The FAA, with the assistance of National Aeronautics and Space Administration (NASA) 

aircraft operations, performed a series of ground and flight tests designed to study the simplified 

inerting concept developed by the FAA. The FAA developed a prototype inerting system, based 

on ASM technology, designed to maintain an inert ullage in the CWT of a 747 commercial 

transport during normal ground and flight operations using a dual-flow methodology. The 

inerting system was mounted in the pack bay of a NASA 747 Space Shuttle Orbiter Carrier 

Aircraft (SCA) to perform the FAA-managed test plan. The system was interfaced with the 

aircraft systems, and instrumentation was installed to measure the inerting system performance 

and fuel tank flammability during the flight tests. Both the CWT and the number 2 wing fuel 

tank (here after referred to as the wing tank) were instrumented with gas sample tubing and 

thermocouples to allow for real-time monitoring of total hydrocarbon concentration, 

temperatures, and CWT ullage oxygen concentration distribution during each test flight. 

2. TEST ARTICLE AND PROCEDURES. 

2.1 TEST ARTICLE. 

The test article consisted of a NASA 747 SCA associated instrumentation and data acquisition 

system (DAS). The FAA OBIGGS with associated instrumentation was installed in the pack 

bay. The CWT was instrumented with eight gas sample tubes that were routed to the FAA’s 

OBOAS, a heated gas sample port routed to the FAA Flammability Analysis System (FAS), as 

well as multiple temperature probes. The wing tank was instrumented with both a heated gas 

sample port for the FAS and multiple temperature probes as well. 

2.1.1 Aircraft. 

The test article was a 747-100, highly modified by NASA, as an SCA (figure 1). It has a basic 

operating weight of 318,053 lbs with a gross taxi weight of 713,000 lbs. It has a 195′ 8′′ wing 

span and is 231′ 10′′ long. The maximum cruise speed of the aircraft is 250 knots or Mach 0.6 

with a ceiling of 35,000 feet. 
FIGURE 1. NASA 747 SCA AIRCRAFT WITH SPACE SHUTTLE ORBITER 

Figure 2 shows the four main wing fuel tanks, the two wing tip reserve fuel tanks, and the large 

CWT of the NASA 747 SCA. The CWT has a capacity of 12,890 gallons of fuel and is between 

the wings of the aircraft, within the fuselage of the aircraft (center wing box). The CWT is 

approximately 242 inches long and 255 inches wide with a height varying from 78 inches to 48 

inches and is partitioned into six bays, two bays are the full width of the fuselage, while another 

two full-length bays are bisected mid-way with a partial rib creating four bays. The empty CWT 

ullage volume was determined to be 1775 normal cubic feet for the purpose of inerting the tank 

with NEA. A large dry bay exists in the center wing box forward of the first tank bay. 

Immediately below the CWT, in an area known as the pack bay, are three air cycle machines 

(ACM) that provide conditioned air to the aircraft and rejects heat to the CWT. 

No. 2 Forward

Main 

Tank 

Center 

Wing 

Tank 

No. 3 

Main 

No. 1 Tank

Main 

Tank 

No. 4 

Main 

Tank No. 4 

Reserve 

Tank 

No. 1 

Reserve 

Tank 

Surge

Tank

Surge 

Tank 

Dry 

Bay 

Dry 

Bay 

FIGURE 2. NASA 747 SCA FUEL TANK DIAGRAM 

3
The CWT is vented to both wing tip surge/vent tanks via a vent channel in each wing that is

plumbed to vent tubing contained within the tank. One vent channel travels along the top of bays 

3 and 4, while the other travels along the top of bays 5 and 6. These channels vent crosswise to 

the exterior of the tank so that the vent channel plumbed on the right side of the tank (bays 5 and 

6) is vented to the left wing surge/vent tank and vise-versa. Each vent channel is plumbed to a 

length of aluminum tubing on each side of the tank that travels forward perpendicular to the 

spanwise beams and midspar across the bays. A smaller tube travels aft within the vent channel 

bay. This plumbing configuration allows the CWT to vent pressure in various rolling and

climb/dive scenarios. The vent channels from the CWT and all other fuel tanks terminate in one 

or both of the surge/vent tanks located near each wing tip. These surge/vent tanks catch fuel and 

prevent overflow, and are connected to the aircraft exterior via a NACA scoop located on the 

bottom surface of each wing. 

2.1.1.1 Replacement of Ventilation Port With Instrumentation Panel. 

The 747 has a ventilation port on the rear spar to ventilate the CWT with fresh air during 

maintenance operations. The FAA installed an instrumentation panel in place of the CWT 

ventilation purge door on the rear spar as shown in figure 3, allowing easy passage of

instrumentation to and from the CWT. The instrumentation panel accommodated ten bulkhead 

fittings for the gas sample lines as well as two thermocouple bulkhead ports for sealing the 

thermocouples at the rear spar. There was also a 1-inch bulkhead fitting in the center of the 

panel for the NEA deposit line. The installation drawing for the instrumentation panel is shown 

in appendix A. 

FIGURE 3. MODIFIED REAR SPAR PURGE DOOR 

2.1.1.2 Modification to the Vent System. 

As described above, the 747 aircraft vents the CWT to both wing tip surge (overflow) tanks. 

Cross venting is the process by which subtle pressure changes between these vents allow outside 

air to ventilate the ullage by passing through one vent channel, into the tank, and out of the other 

vent channel. 

4
5

To prevent the loss of inert gas by this process, the FAA installed three blocking plates in the 

right side vent system (looking forward) of the NASA 747 SCA. One was installed by removing 

the vent flange that attached the vent tubing to the vent channel, placing a thin aluminum plate 

between the flange and the vent stringer and sealing it with fuel tank sealant. In addition, the 4-

inch-diameter vent tube also had a thin, round aluminum plate installed in one connection fitting 

to prevent inert gas from transferring through the tube from one bay to another. It was not 

necessary to use fuel tank sealant on this blocking plate. The right side vent stringer also has a 

fuel drain with a float valve mounted on it. The float valve was removed and a blocking plate 

was installed in its place and sealed with fuel tank sealant. This modification prevents cross 

venting of the CWT ullage. 

After the FAA completed the tests to examine inerting without cross venting, the vent system 

was restored to the original equipment configuration for a single baseline flammability test. 

2.1.1.3 Gas Sample Fittings. 

Two sample port tubes were installed through modified HiLok fasteners, which were installed 

in place of specified existing fasteners in the wing structure, to allow for the measurement of 

total hydrocarbons in both the CWT and the wing fuel tank. In each case, an appropriate fastener 

was specified, modified with a 1/8-inch hole through the fastener center, and then installed in the 

appropriate location. The CWT fitting penetrated the top of the tank in the number 2 bay slightly 

left of centerline, while the wing tank fitting was located just outboard of the number 2 engine 

pylon on the front spar. The sample port tubes were later bonded into the fastener holes with 

aviation grade epoxy, and float valve assemblies were installed on the tube inside the tank to 

prevent sloshing fuel from entering the sample lines attached to the tubing outside of the tank. A 

drawing illustrating the hollow-fastener installation with float valve assemblies is shown in 

appendix B. There was a third tooling hole fastener modified on the left wing forward spar to 

accommodate the eight thermocouples installed in the number 2 main fuel tank. When 

completed, the installation was sealed with the appropriate fuel tank sealant. 

2.1.2 Instrumentation.

The primary instrumentation centered around the heated and unheated gas-sampling tubing and 

thermocouples in both the CWT and the wing tank. Continuous measurement of oxygen 

concentration and total hydrocarbon concentration (THC) in the fuel tanks during each flight test 

were made using the FAA OBOAS and FAS. Additionally, the FAA OBIGGS was instrumented 

with pressure, temperature, oxygen concentration ports, and a flow sensor to analyze system 

performance and health monitoring for each test. Additional test parameters were also collected 

to aide in the analysis of the inerting process and progression of flammability during the flight 

tests. Appendix C contains a complete instrumentation diagram as well as a complete list of all 

instrumentation channels. 

2.1.2.1 Onboard Oxygen Analysis System.

The FAA OBOAS was used to measure the oxygen concentrations at the eight specified 

locations within the CWT. The OBOAS was developed by the FAA to measure the oxygen 

concentrations in a fuel tank environment during a typical commercial-transport airplane flight 
6

cycle using conventional oxygen analyzers. This system consists of a regulated sample train 

with flow-through, in-line oxygen sensors and ancillary equipment. Two identical four-channel 

systems were developed. Each four-channel system was self-contained in a 19-inch flight test 

half-rack. Each system has four independent sample trains that can draw an ullage sample at 

four different locations in the fuel tank, regulate the sample pressure, expose the sample to the 

oxygen sensor, and redeposit the sample back into the fuel tank. Each oxygen sensor has a 

companion analyzer mounted on the same 19-inch half-rack. Also mounted on each 19-inch 

half-rack is a four-channel inlet pressure controller and a single-outlet pressure controller 

electronic unit. These electronic units support the five pressure regulator/controllers in each 

four-channel system. Reference 7 gives a complete description of the measurement system with 

part lists, diagrams, and systems analysis. 

2.1.2.2 Flammability Analysis System.

The FAA developed the FAS for the purpose of real-time, in-flight monitoring of the THC in 

both the CWT and wing tank. The FAS consists of a two-channel Rosemount Analytical NDIR 

analyzer, supplied by temperature-controlled, pressure- and flow-regulated sample streams. The 

NDIR analyzer was custom-designed and constructed by the manufacturer to safely measure fuel 

tank vapors in flight and housed in an explosion proof box. 

All hydrocarbon gas samples must be maintained at a minimum 200°F temperature prior to 

entering the NDIR analyzer to eliminate any condensation of the hydrocarbon vapors prior to 

being measured. To that end, the sample stream external to the FAS consists of several heated 

sample tubes and a heated rack-mounted box, which houses the diaphragm pump heads, flash 

arrestors, float valves, sample control valves, and sample flow meters. A separate unheated, 

rack-mounted box houses the two inlet sample pressure regulators. For safety, both rackmounted boxes are continuously purged with air in the same manner as the OBOAS to prevent 

the accumulation of explosive vapors in the event of a sample line leak. An additional rackmounted panel houses all necessary electronics to maintain the proper sample conditions, 

including temperature controllers and the pressure controller electronic units. 

Figure 4(a) shows a diagram of the rack-mounted installation of the pump, sample-conditioning 

boxes, and controller box, and figure 4(b) shows the NDIR analyzer. 

2.1.2.2.1 Sample Train Flow.

After entering the FAS heated rack-mounted box, the samples pass through a flash arrestor and 

then a four-way selector valve. The selector valve enables the operator to draft a sample from 

the tank, cabin air, calibration gases, or to close off the sample inlet. Each sample then passes 

through a two-stage diaphragm pump. Prior to exiting the heated rack-mounted box, the samples 

are teed off to the inlet pressure regulators located in the sample-conditioning unit, which 

regulates the inlet pressure. On the other side of the tee, the samples pass through a sample flow 

meter prior to exiting the heated rack-mounted box. Another heated line then routes the samples 

to the NDIR analyzer. On the exit side of the NDIR analyzer, the two samples are joined 

together and dumped overboard after flowing through a manual backpressure regulator. 
Heated Sample

Box 

Sample

Flow/Pressure

Conditioning Unit

Controller 

Electronics Panel 

 

(a) (b)

FIGURE 4. (a) RACK-MOUNTED FAS INSTALLATION AND (b) NDIR ANALYZER 

In the event that a float valve is actuated and flow through that line is blocked, the operator has 

the capability to purge that line once it drops below the fuel level. This is accomplished by

sampling from the cabin and circulating that sample back through the closed portion of the line 

until the float opens back up. A schematic of the sample train flow is shown in figure 5. 

FIGURE 5. THE FAS SAMPLE STREAM FLOW SCHEMATIC 

2.1.2.2.2 System Safety Features.

Careful measures were taken to ensure that the FAS operates in such a manner to minimize the 

operator’s and crew’s exposure to hazardous vapors as well as to preclude the ignition of any 

potentially flammable vapors that may be drafted through the FAS. 

The NDIR analyzer is housed within a continuously purged housing with an impact-tested, 

intrinsically safe front panel. The electronics of the NDIR analyzer have been separated from the 

7
sample stream to allow safe heating of all internal sample tubes to 200°F as well as to separate

any potential ignition sources from the sample stream in the event of a leak of flammable vapors.

The purge system incorporated an easily observed indicator of a safe or nonsafe condition in the

form of a rotary flow indicator. If purge pressure was unable to maintain a safe condition, the 

unit was immediately shutdown. The purge system for the rack-mounted sample-conditioning

boxes used two ejectors each requiring a minimum of 3 standard cubic feet per minute (SCFM) 

of compressed air at 40 psi or greater. The ejector unit maintained a constant negative pressure 

on the inlet, drafting a volume of air from inside the boxes. The ejector inlet is plumbed through 

a rotary flow indicator, which allowed the operator to determine if the purge flow is functional at 

any point in time. 

Likewise, the rack-mounted sample-conditioning boxes, housing the pump, flow meter, and 

sample pressure regulators, are continuously purged enclosures and are separated from the rackmounted panel housing the control electronics (temperature controllers and pressure control 

electronic units). In addition, care has been taken to only mount the diaphragm pump head 

inside the heated box. The motor driving the diaphragm pump was kept external to the box, 

again to maintain separation between any electrical sources and the sample stream flow. 

2.1.2.3 Gas-Sampling Lines. 

Eight sample ports are located in the CWT at the eight locations identified in figure 6. The 

sample lines are made of PFA tubing routed from bay 6 to each location through the fuel seep

holes at the top of the tank between the spars/spanwise beams and the side of body ribs. The 

tubing is terminated at a float valve that is attached to the tank through a mounting plate that was 

attached to a stiffening bracket between two stringers (with a stiffener) at the top of the tank. 

This small plate is attached to the stiffener bracket through a replaced rivet to minimize the 

necessary modification for the installation. Figure 7 shows one sample port mounting plate 

attached in the CWT with a float valve assembly. 

FIGURE 6. TOP VIEW OF A 747-100 CWT WITH SAMPLE PORT LOCATIONS 

8
FIGURE 7. GAS SAMPLE FLOAT VALVE INSTALLATION 

2.1.2.4 Heated Sample Lines. 

The fuel vapor samples were drafted from the fuel tanks via the installed sample fittings (see

section 2.1.1.3). After exiting the tanks, the samples were routed to the FAS using a series of 

heated sample lines maintained at a constant 200°F. For the CWT sample, a single 10-foot 

heated line was routed directly from a modified fastener at the top of the tank to the rackmounted portion of the FAS. The sample from the wing tank was routed from a similarly 

modified fastener on the front spar, down the leading edge of the left aircraft wing with a 50-foot 

heated line. This line was routed to a fitting on the fuselage pressure bulkhead with a separate 

heated line routed from this fitting, into the front cargo bay, and to the FAS. The temperature of 

all heated lines along with the heated box was continuously monitored by the operator via panelmounted displays. In addition, the 50-foot line was constructed to have additional 

thermocouples located every 5 feet along the line that were recorded by the DAS. 

2.1.2.5 Fuel Tank Thermocouples. 

There were 24 thermocouples mounted in the CWT positioned at the locations identified in

figure 8. The T-type thermocouples were routed through compression fittings on the instrument

panel and sealed with fuel tank sealant. Fourteen of these thermocouples were mounted on 

metallic surfaces with an epoxy-retaining patch (appendix D), while the remainder were

suspended at the desired location. The thermocouples were mounted to the tank structure using 

fuel-compatible wire ties with some long runs being supported by additional epoxy patches. The 

thermocouples were 1/16-inch stainless steel sheathed and ran from the panel installed in bay 6 

to each desired location through the fuel seep holes at the top and bottom of the tank between the 

spars/spanwise beams and the side of the body ribs. 

9
Side of

Body Rib

Side of 

Body Rib 

FIGURE 8. TOP VIEW OF A 747-100 CWT WITH THERMOCOUPLE LOCATIONS 

The ullage thermocouples (1 through 8) were installed within 1 foot of the bottom of the caps of 

the ceiling stringers. The fuel thermocouples (21 and 22) were installed less than 1 inch from the

tank bottom. Thermocouples 9 through 16 were mounted on the floor. Thermocouples 17-20 

were mounted on the tank walls approximately half way between the bottom and the top of the

tank. Thermocouples 23 and 24 were mounted on the ceiling of the tank. 

There were eight thermocouples mounted in the wing tank, positioned as described in table 1.

The thermocouples were the same type used in the CWT and were routed from the cabin, 

through the cargo bay, into the fairing area, along the leading edge, and through a drilled out 

tooling hole on the front spar of the wing approximately at wing station (WS) 400. Five of these 

thermocouples were mounted on surfaces in the same manner as in the CWT, with the remainder 

suspended at the desired location. Figure 9 gives a top diagram of wing tank number 2 with the 

approximate location of the thermocouples noted. 

10
TABLE 1. WING TANK THERMOCOUPLE LOCATIONS 

Temp

No. Description Location

1 Wing tank forward spar surface 

temperature 

Front spar surface, WS 455, 25′′ from tank 

bottom

2 Wing tank inboard fuel 

temperature 

4′ from front spar, WS 310, 6′′ from tank 

bottom

3 Wing tank rear spar surface 

temperature 

Rear spar surface, WS 435, 25′′ from tank 

bottom

4 Wing tank ullage temperature 2′′ from front spar, WS 650, 4′′ from tank top 

5 Wing tank mid-fuel/ullage

temperature 

25′′ from front spar, WS 586, 13" from tank 

bottom

6 Wing tank outboard wall surface 

temperature 

6′ from front spar, WS 715, 12′′ from tank 

bottom

7 Wing tank bottom surface 

temperature 

8′ from front spar, WS 455, tank bottom

surface 

8 Wing tank top surface temperature 8′ from front spar, WS 455, tank top surface 

Ullage Gas Sample

Location

1 

2 

4 

3

5

6 

7 8

WS 400 

WS 630 

WS 455

WS 586

WS 499 

Thermocouple

Penetration

Front Spar

#2 Engine

Pylon

FIGURE 9. TOP VIEW OF A 747-100 WING TANK WITH 

THERMOCOUPLE LOCATIONS 

11
12

2.1.2.6 Onboard Inert Gas Generation System Parameters. 

The FAA OBIGGS was equipped with instrumentation that monitored and set system 

parameters. The instrumentation was routed from the OBIGGS mounted in the empty right side, 

rear pack bay area to the cabin. The inerting system was instrumented with six thermocouples 

and four pressure taps to monitor system health and performance. Pressure transducers, used to 

monitor the system pressures, were mounted on a panel in the cabin with pressure-sensing tubes 

routed from the pressure taps to the transducer panel. All OBIGGS pressure tap and temperature 

measurement locations are identified on the OBIGGS drawing given in appendix E. 

Additionally, ambient pressure was measured in the pack bay for data reduction. 

The OBIGGS was also equipped with gas-sampling ports to measure both NEA and oxygenenriched air (OEA) oxygen concentrations, also shown on the drawing in appendix E. These 

oxygen samples were monitored by a two-channel oxygen analyzer capable of sampling from 

altitude, which operated very similar to the OBOAS. This oxygen analyzer did not require fluid 

traps, flash arrestors, or containment box purging because the sample system did not acquire a 

sample from a potentially flammable source. Also, a flow meter was installed at the NEA output 

to measure the NEA flow rate. These parameters constituted the ASM performance at a given 

flight condition. 

2.1.2.7 Additional Parameters Collected. 

Additional instrumentation was employed to determine the validity of certain data and to better 

understand the fuel tank inerting process, in general, as it relates to the FAA concept. Static 

pressure was measured, allowing examination of the effect of altitude on the inerting process. 

The NEA temperature at the flow meter and the backpressure on the OEA line was also acquired. 

Thermocouples were also located in the pack bay (4), one of which was in the immediate area of 

the OBIGGS. An additional three thermocouples were installed in the area of the bleed air 

connection to the system. This allowed for redundant measurement of temperature in the area of 

the pack bay that could potentially be susceptible to a bleed air leak.

2.1.2.8 Data Acquisition System. 

A DAS recorded all analog signals required for the test in a multiplexed method and stored the 

data on the flight test data acquisition computer. Some data were presented on a master display, 

in a block diagram format, to facilitate real-time monitoring of the more significant parameters 

during the testing. Figure 10 shows the master DAS display. 

2.1.3 Aircraft Cabin Installation. 

All instrumentation was routed to the cabin via several different cabin penetrations to an array of 

FAA-designed and built instrumentation racks installed in the central part of the pressurized 

cabin immediately above the CWT. All instruments, as well as the OBIGGS, were powered via 

an extensive power distribution panel wired to an existing work rack immediately forward of the 

FAA rack system. 
FIGURE 10. MASTER DAS DISPLAY 

2.1.3.1 Power Distribution. 

All power required to run the OBIGGS and the test instrumentation was obtained from an 

extensive power distribution bus located on an existing utility rack at station (STA) 990. The 

OBIGGS required both 115 Vac, three-phase 400 Hz and 28 Vdc power to operate. Each rack

(2) in the OBOAS required 115 Vac, three-phase 400 Hz to operate an accompanying 115 Vac, 

single-phase 60-Hz power voltage converter rated for 20 amps. The two voltage converters are 

used to power the OBOAS as well as the OBIGGS oxygen analysis system with a power 

switching system attached to the OBOAS rack system. 

An existing 60-cycle power converter was used in conjunction with an additional NASA power 

converter to power the FAS with the associated instrumentation. This included the NDIR 

analyzer, all heated lines, and pressure controllers. Additionally, the computer and DAS were 

powered by this set of power converters using an uninterruptible power supply. Figure 11 

illustrates the distribution of power for the flight test instrumentation and DAS. 

2.1.3.2 Instrumentation Routing. 

The CWT thermocouples were routed from the modified rear spar purge door into the cabin 

through an existing wiring penetration in the right-hand wheel well at STA 1275. The DAS was

located on the same rack with the OBIGGS control box and gas analysis equipment. The wing 

tank thermocouples were routed from a modified fastener in a tooling hole on the front spar, 

along the front spar into the fairing area, and into the fuselage through another wiring penetration 

13
at STA 985. These thermocouples were routed to the DAS rack under the floor in the cargo bay 

to the right side of the cabin. 

OBIGGS

Analysis Pump

OBOAS

Pump 1

OBOAS

Pump 2

OBOAS Rack Assembly

OBIGGS

Control

Box

DAS

Computer

DAS &

OBIGGS

Rack

NDIR

Analyzer

Sample

System

FAS 

Rack

Power Strip UPS

Ad 115 VAC, 400 Hz Bus ditional

60 Hz Power

Converter

28 VDC Bus

Existing Power Rack

OBIGGS Ansy

Sample Train

Existing

60 Hz Power

Converter

Sample

Pump

FIGURE 11. POWER DISTRIBUTION FOR THE NASA 747 SCA CABIN 

The eight CWT gas-sample lines and two sample return lines were routed from the rear spar 

purge door fittings to the OBOAS in the cabin via a cabin penetration in the wheel well adjacent 

to the thermocouple penetration (approximate STA 1250). The CWT gas-sample lines were 

routed into the cabin through a mating fixture that terminated the shrouds of the eight sample

lines and two returns in the cabin area. The lines were routed to the OBOAS location in the

center of the aircraft adjacent to the DAS rack aft of STA 1080. Each four-channel system had a

gas sample return routed to the tank through the same mating fixture to minimize the effect of 

the sample system on the ullage environment. The signals from the eight oxygen analyzers were 

routed to the DAS. 

For the CWT THC gas sample, a single 10-foot heated line was routed directly from a modified 

fastener in the cabin near STA 1100 to the FAS rack located on the left side of the cabin near 

STA 1030. The sample from the wing tank was routed from a similarly modified fastener at 

WS 630 on the front spar to a 50-foot heated line routed down the leading edge of the left aircraft

wing, through an existing hole between the leading edge and the fairing area, to a bulkhead 

fitting on the fuselage pressure bulkhead in the front cargo bay at STA 985. From this fitting,

14
another 10-foot heated line conveyed the sample to the rack-mounted portion of the FAS. The 

FAS also required two heated lines between the FAS rack and the NDIR analyzer immediately 

adjacent to the rack at approximately the same fuselage station. Signals from the FAS were

routed across the cabin to the DAS rack. 

The control cable for the OBIGGS was routed from the OBIGGS in the pack bay to the OBIGGS 

control box through a modified waste shoot that provided a conduit to the cabin from the aft

bulkhead of the pack bay (figure 12). All OBIGGS instrumentation, including gas-sampling 

tubing, pressure-sensing tubing, and thermocouples, as well as all pack bay thermocouples were 

also routed in the same manner. The waste shoot was capped at the floor in the cabin area with 

an aluminum plate sealed in place, which had a slot to accommodate the instrumentation bundle. 

Additional sealant was used to seal the slot around the instrumentation bundle. 

FIGURE 12. MODIFIED WASTE SHOOT USED AS A CONDUIT AT THE AFT PACK 

BAY AREA BULKHEAD 

Figure 13 illustrates the cabin layout and equipment proximity as well as basic signal routing for 

the flight test setup. 

15
Aircraft Cabin

FAS

Air

OBOAS

Existin

g

Power

Rack

Window Plug

DAS

OBIGGS

Air

Air

Air

OBOAS

Compressed Air Bank

(Cargo Bay)

Penetration STA 1275

(CWT Thermocouples)

Penetration STA 1235

(Pack Bay Instrumentation)

Penetration STA 985

(Wing Tank Instrumentation)

Power Cable

System Wiring

Sample Tubing

Purge Air Line

Instrumentation

Wiring

To Avionics Bay

STA 990 STA 1010 STA 1080 STA 1120 STA 1250

FAS

Analyzer Sample System

Penetration STA 1244

(CWT Sample Lines)

Wing Tank/Heated

Line Thermocouples

FIGURE 13. DIAGRAM OF INSTRUMENTATION INSTALLATION IN THE 

NASA 747 SCA CABIN 

2.2 TEST PROCEDURES.

Nine flight tests were performed for a total of 29 hours. An additional 11 hours of ground testing 

was performed prior to and during the execution of the flight tests. All tests were performed 

from Ellington Field, TX, the center of NASA aircraft operations, with the exception of test 0, 

which departed from NASA Dryden, Edwards, CA. The ground operation between tests 4 and 5 

occurred at Davis-Mothem Airbase in Tucson, AZ. The tests were performed during two 

different 2-week periods approximately 4 months apart. The primary focus of the testing was to

study commercial transport fuel tank inerting using the FAA-simplified inerting system. Both 

the dual-flow and variable-flow operation was studied, with and without fuel in the CWT. 

Ullage flammability was measured under a variety of flight and ground conditions in both the 

CWT as well as in the wing tank. Additionally, flammability was measured during tests without 

inerting and with a conventional CWT vent configuration. Table 2 gives a list of flight tests 

performed on the NASA 747 SCA during the fuel tank inerting and flammability study. 

16
17

TABLE 2. TABLE OF INERTING FLIGHT TESTS ON THE NASA 747 SCA 

No. 

FAA 

Designator Date 

Flight 

Hours Description 

1 Pretest 1 12/09/03 2.1 Preliminary OBIGGS proof of concept. CWT approx. 

11% [O2], run system gate to gate, ascend to FL 350, 

cruise until CWT [O2] stable, descend per ATC. CWT no 

cross venting. 

2 Pretest 8 12/10/03 2.9 OBIGGS performance evaluation. Stabilized system 

inputs at three altitudes for three different flows and two 

different temperatures. CWT no cross venting. 

3 Test 0 5/13/04 3.8 FAS Evaluation. CWT not inert, OBIGGS off, operate 

FAS gate to gate. Operate ACMs on ground to generate 

flammability. CWT no cross venting. 

4 Test 1 5/19/04 2.1 OBIGGS Demonstration–Dual Flow Mode. CWT 

approximate 10% [O2], run system gate to gate, ascend to 

FL 350, cruise until CWT [O2] stable, descend per ATC. 

CWT no cross venting. 

5 Test 2 5/21/04 2.1 OBIGGS Variation 1–Variable Flow (small orifice). 

CWT approx. 10% [O2], run system gate to gate, ascend 

to FL 350, Cruise until CWT [O2] stable, descend same as 

test 1 within limits of ATC. CWT no cross venting. 

6 Test 3 5/22/04 2.1 OBIGGS Variation 2–Maximum Flow (wide orifice). 

CWT approximate 10% [O2], run system gate to gate, 

ascend to FL 350, cruise until CWT [O2] stable, descend 

same as Test 1 within limits of ATC. CWT no cross 

venting. 

7 Test 4 5/20/04 4.2 Flammability Reduction Demonstration Part 1–OBIGGS 

Variable Flow. CWT approximate 10% [O2], 25% fuel 

load, run system gate to gate, ascend to max altitude, 

cruise until CWT fuel burned and [O2] stable, descend per 

ATC, 3-hour turn-around running ACMs, CWT no cross 

venting. 

8 Test 5 5/20/04 3.6 Flammability Reduction Demonstration Part 2–OBIGGS 

Variable Flow. CWT approximate 10% [O2], 25% fuel 

load, run system gate to gate, ascend to max altitude, 

cruise until CWT fuel burned and [O2] stable, descend per 

ATC, 3-hour turn-around running ACMs, CWT no cross 

venting. 

9 Test 6 5/26/04 6.1 FAS Evaluation. CWT not inert, OBIGGS off, operate 

FAS gate to gate, operate ACMs on ground to generate 

flammability. CWT venting original configuration. 
3. ONBOARD INERT GAS GENERATION SYSTEM.

The FAA prototype OBIGGS was designed to incorporate a dual-flow concept, developed by the 

FAA, that provides a low, relatively pure inert gas flow to the tank during taxi, takeoff, and 

cruise to obtain a low oxygen concentration. During descent, the system was switched to lowflow mode, which provides a low amount of less pure NEA to the tank to reduce the amount of 

air entering the tank through the vent system. The system was modified to allow for a 

variable/high flow during the system operation. This allowed the operator on some tests to 

increase the amount of flow at higher altitudes to decrease the total amount of air entering the 

ullage through the fuel tank vent system during descent. 

The FAA OBIGGS consists of a single unregulated flow path that is plumbed to a manifold of

three ASMs. The nitrogen-rich gas that passes through the modules is then plumbed to the CWT 

to reduce the oxygen concentration. The flow path has a heat exchanger that controls the ASM 

inlet temperature and a filter. The flow passes through the ASMs and through the flow control 

valves, which allows the system to flow in both low- and variable/high-flow modes. Figure 14 

shows a block diagram of the primary components of the OBIGGS. The system is mounted on a 

relatively simple aluminum frame in a palletized manner for simplicity of assembly and 

installation.

Heat

Exchanger

Cool Air Source

Existin

g Aircraft Bleed

air Sup

ply

Filter

Heater

Shut-Off

Valve

Variable Flow

Valve

Fuel Tan

k

Throttling

Valve

AS

M

System Controller

Temperature

Sensor

AS

M

AS

M

Aircraft

Po

wer

Oxygen Rich

Waste Gas

Waste Stream

Discarded

Nitrogen Rich

Product Gas

Low Flow

Orifice

FIGURE 14. SYSTEM BLOCK DIAGRAM 

3.1 SYSTEM FLOW.

The inerting system gets its bleed air from a modified main bleed air duct on the aft bulkhead of 

the 747 pack bay at STA 1245, which is reduced to a 2-inch-diameter tube for connection to the 

18
19

system shutoff valve (SOV). When energized, the 18-32 Vdc SOV valve opens, allowing flow 

into the system, through the heat exchanger bypass Y, into the air/air heat exchanger. The heat 

exchanger accepts the 400°F bleed air and uses a 4-inch-diameter cooling bypass loop to cool the 

system flow to a temperature of 180°F ±10°. The cooling bypass uses a three-phase, 115 Vac, 

400-Hz fan to draw outside ambient air through the heat exchanger and uses a 4-inch-diameter, 

18-32 Vdc motor-operated, modulating valve to control the airflow through the heat exchanger. 

This modulating valve is operated by the system temperature controller, but is controlled 

manually with a switch box (referred to as the Parker box) by the system operator. The heat 

exchanger bypass Y allows a 1-inch-diameter line to bypass a portion of the system flow into the 

heat exchanger to decrease the effectiveness of the cooling loop, giving better control of the air 

temperature into the ASMs. 

After the heat exchanger, the bleed air passes through a desiccating filter, past the temperature 

sensor used by the system operator to control the temperature, and through a section of pipe with 

a 187-watt clamp heater before entering the ASMs via a manifold. The clamp heater is designed 

to increase the temperature of the ASM inlet air significantly if the temperature should drop (due 

to system heat rejection) well below the target temperature obtained by the heat exchanger. 

However, it provided very little additional temperature when the air was maintained close to the 

specified temperature. 

Once the 180°F cooled bleed air passes into the ASMs, the air is separated into the NEA and 

OEA constituents. The NEA portion passes through to the ASM outlet, while the OEA portion 

passes out the permeate (waste) port. The ASM waste flow is eliminated from the system 

through the OEA manifold plumbed into the heat exchanger exhaust box on the bottom side of 

the aircraft. 

After the OEA is separated, the NEA passes through the flow control portion of the system, 

which allows the system to flow at either low-flow (low-oxygen concentration), or variable/high 

(higher-oxygen concentration) flow conditions. The variable-flow valve can be closed forcing 

all the flow through the low-flow orifice, or it can be opened to one of five distinct open 

conditions, of different backpressure, allowing the system to give a total of six different flow 

conditions. As the variable-flow valve is opened, the system backpressure will decrease, 

creating progressively more NEA flow, with the NEA oxygen concentration increasing (less 

pure) as flow increases. Two separate variable-flow valves were used; both were 1-inch, motoroperated, ball-type SOVs. One had a standard ball with a 1-inch-diameter hole, the other had a 

square slot with an effective 0.391-inch-diameter hole. 

The NEA can be directed by means of counter-actuated SOVs to either “inert” the CWT or to 

“divert” the NEA overboard. When divert is selected, the NEA flows directly to the heat 

exchanger exhaust box where it mixes back with the OEA and is safely dumped overboard. If 

inert is selected, the NEA flows through a flow meter, a check valve, and finally to the NEA 

deposit line. 
3.2 SYSTEM INTERFACES.

The primary inerting system physical interfaces are the attachment brackets that mount the 

system to the aircraft. The mechanical interfaces allow bleed air to interface with the system, 

NEA to be deposited into the CWT, OEA to be eliminated, and heat exchanger air to be used.

Additionally, system electrical connections also provide an interface between the system and the 

aircraft power as well as between the system and the operator. The system diagram given as

appendix E identifies each system interface. 

3.2.1 System Mounting.

To allow a relatively simple interface with the aircraft, the system pallet was installed in the

number 4 pack bay of the NASA 747 SCA, frequently referred to as the empty pack bay, using 

six attachment brackets. The brackets attached the aluminum angle pallet to the fairing support 

structure in the aft right-hand side of the pack bay area. The brackets were made of stainless

steel and were attached using doublers with 1/4′′ Steel A/N standard bolts. Figure 15 shows a 

three-dimensional rendering of the system mounted in the empty pack bay of a 747 with the main 

bleed air duct illustrated in blue as a spatial reference. 

FIGURE 15. THREE-DIMENSIONAL RENDERING OF THE OBIGGS 

INSTALLED IN THE PACK BAY 

3.2.2 Mechanical Interface.

The primary mechanical interfaces for system operation were the bleed air inlet, the NEA 

deposit, the OEA discharge, and the heat exchanger cooling air inlet and exit. Bleed air was

supplied to the system by modifying the bleed air duct (Boeing part no. 69B41284-12) at STA 

1245. The modification consisted of welding a 2-inch, commercially pure titanium BMS7-21 

grade 3 tee into the main bleed air duct, and using a sealed sleeve assembly to connect the bleed 

air duct to the OBIGGS SOV assembly. The system SOV was a 2-inch-diameter, 28-32 Vdc 

20
21

solenoid operated, mechanically closed valve, Parker part number 5940016 that had a 55-psig 

overpressure protection circuit. The overpressure protection feature closes the valve if the static 

pressure on the system exceeds 55 psi, protecting the system components in the event of a bleed 

air system over-pressurization. 

The OBIGGS NEA output was a 1-inch HydroflowTM 14J02 sleeve and coupling connection. 

The NEA inert/divert tubing assembly, described in section 3.1, was a series of 1-inch 

HydroflowTM fittings with associated 1-inch SOVs adapted to a 16D A/N fitting for mating with 

the NEA deposit (see appendix F). The NEA was carried to the CWT through a 1-inch flexible 

AE701 wire-reinforced hose with stainless steel over braid lines using MIL-F-83798 specified 

816-16D standard fitting ends. The 1-inch flexible line attached to an AN837-16D bulkhead 

fitting on the outboard, aft pack bay-canted bulkhead, which passes the NEA into the wheel well 

area. A second flexible line passed the NEA from the A/N bulkhead fitting to the rear spar purge 

panel. The NEA entered the CWT through a 1-inch-diameter SwageLok® bulkhead fitting on the 

rear spar purge panel (see figure 3). Adaptors were used to adapt the A/N fittings to the 

SwageLok bulkhead fitting. A third 1-inch flexible line was used to carry NEA to a nozzle 

mounted on a bracket at the top of bay 6 (right aft most bay) in the CWT. This nozzle consisted 

of a 90o elbow 16D A/N bulkhead fitting directing the NEA down, and a check valve was placed 

immediately before the nozzle to reduce the possibility of fuel seeping back into the inerting 

system. Appendix G contains an installation drawing of the NEA nozzle. 

The OEA discharge was connected to the inboard side of the heat exchanger exhaust box via a 2-

inch inner diameter (ID) aluminum tube and flexible coupling. A 2-inch check valve was also 

used to keep contaminants from entering the ASMs when not in operation. 

The heat exchanger cooling air was interfaced with the pack bay exterior through two scoops on 

two separate fairing panels that were modified for the flight tests. The heat exchanger inlet door 

was located on fairing panel 192KR, between STA 1169 and 1181, and was approximately 10.5 

inches above the bottom edge of the panel. The air inlet door had a semi-circular cross section or 

approximately 6 inches in diameter and was mounted in a 13.5- by 12- by 7-inch assembly. It 

was connected to a 4-inch motor-operated air inlet modulating valve using two flexible couplings 

and an aluminum tube, which had a 2-inch opening into the air inlet box. The heat exchanger 

cooling air exhaust door had a rectangular cross section approximately 6 inches by 3 inches and 

was located on access door 192NR, between STA 1223 and 1233, approximately 10.5-inches 

from the seam of panel 192KR. The cooling air exhaust door was mounted in an assembly 10.5 

by 8 by 16 inches and was connected to the heat exchanger by a 4-inch flexible coupling that 

attached to an aluminum box fitted to the back of the exhaust door assembly. The NEA bypass 

line was connected with a 1-inch ID stainless steel flexible hose to a 16D AN fitting on this box. 

Figure 16 shows the OBIGGS installed in the empty pack bay of the NASA 747 SCA test 

aircraft without the two panels described above. 
FIGURE 16. ONBOARD INERT GAS GENERATION SYSTEM INSTALLED IN 

TEST AIRCRAFT PACK BAY 

3.2.3 Electrical Interface.

Each system part had a specified electrical connector to communicate power and signals to the 

individual component. They were wired with aviation-grade wiring to a system connector, 

allowing the system to easily interface to the OBIGGS control box. The control box allowed the 

system and its components to be turned on or off and to switch the system from one flow mode 

to the other (see figure 17). It also provided the system operator with the ASM inlet temperature 

and a manual temperature control box. A complete wiring diagram of the system and control 

box is shown in appendix H. 

FIGURE 17. FRONT PANEL OF OBIGGS CONTROL BOX 

3.3 SYSTEM OPERATION.

The system is designed to operate during normal aircraft flight and ground operations, provided 

bleed air is available. The system low-flow and high-flow conditions were established by setting 

two different needle valves during ground operations. The needle valves provide the 

backpressure needed to give the necessary system variable performance throughout the flight 

profile. The aircraft’s auxiliary power unit was used to power the bleed air system to

approximately 30 psi, and the low-flow mode needle valve was adjusted to generate NEA with 5 

percent oxygen at sea level conditions. The ASM inlet temperature was allowed to stabilize at

22
23

180°F before setting the needle valve with the variable/high-flow valve closed. For the first test, 

the high-flow mode needle valve was set to generate NEA with 11 percent oxygen at sea level. 

After the first test, the variable-flow valve was installed, and the high-flow needle valve was 

removed. Instead of a fixed needle valve configuration, the NEA conditions (flow/purity) were 

changed by opening and closing the valve in increments. The low-flow mode was primarily used 

for taxi, takeoff, climb, and cruise phases of flight, while the variable/high-flow mode was used 

primarily for the descent phase of flight and during the ground-taxi back to the gate. 

It is considered essential that the OBIGGS be maintained at 180°F ±10° for efficient 

performance. The system operator controls the ASM inlet temperature manually using the 

system temperature controller. The operator adjusts several potentiometers that control the 

position of the 4-inch-diameter cooling air modulating valve for the heat exchanger. This 

modulating valve needed to be continually adjusted due to constantly changing conditions 

(altitude, bleed air pressure, pack bay temperature) on the aircraft during a typical flight. The 

operator used the temperature readout display on the OBIGGS control box to monitor cooling 

system performance. The readout display received its signal from a T-type thermocouple that 

was installed just before the ASM inlet manifold. 

The operator could use the heat exchanger bypass valve if necessary to bypass a small amount of 

flow (10%-20%) around the heat exchanger to decrease the affectivity of the heat exchanger. 

This made controlling the temperature on the ASMs easier under some conditions. 

During system warmup, maintenance, and testing, there was a need not to effect or disrupt the 

CWT environment, so a NEA divert valve was used to direct the NEA product away from the 

CWT. The inert/divert switch was located on the OBIGGS control panel and operates two 

opposing 1-inch-diameter, 28-32 Vdc motor-operated SOVs (see appendix F). When divert was 

selected, the inert SOV closes, and the divert SOV opens and deposits the NEA overboard. 

When inert was selected, the divert SOV closes, and the inert SOV opens, allowing the NEA to 

travel to the CWT. 

The system required the operator to change from low to variable/high-flow mode manually. 

Low-flow mode gave the most pure (lowest oxygen concentration) NEA the system could 

generate with the fixed low-flow orifice and the given aircraft conditions. Variable/high-flow 

mode was adjusted by incrementing an “Increase/Decrease” switch on the OBIGGS control 

panel. The increase/decrease switch controlled a 1-inch-diameter, 28-32 Vdc motor-operated 

SOV. The variable/high-flow mode can allow the system to produce higher volumes of NEA of 

greater oxygen concentrations (less pure) at higher altitudes. This system feature allows greater 

flow into the tank during aircraft descent to minimize air entry into the ullage due to increasing 

static air pressure. This creates lower resulting ullage oxygen concentration than if the low-flow 

mode was simply employed the entire flight cycle. 

4. ANALYSIS.

Data analysis was performed in two different categories: inerting calculations and flammability 

calculations. 
4.1 INERTING CALCULATIONS. 

The primary calculations performed to analyze the inerting test data were to determine the 

quantity of bleed air consumed by the system and the bulk average oxygen concentration. 

Additionally, a simple inerting model was used to calculate the oxygen concentration in a tank 

volume, given a flight profile and performance schedule of the OBIGGS. 

4.1.1 Bleed Air Consumption. 

System bleed air flow was calculated using an equation developed from flow in and out of the 

ASM and a mole balance of oxygen in and out of the ASM. The combination of these equations 

gives the following equation for bleed air flow in terms of NEA flow. 

(0.21 [O ] )

([O ] [O ] )

2

2 2

Perm

NEA Perm QBleed QNEA −

− = ⋅ & & (1) 

With: = NEA Oxygen Concentration NEA [O ]2

Perm [O ] 2 = OEA Oxygen Concentration 

A more complete derivation of the equation is given in reference 5. 

4.1.2 Average Fuel Tank Oxygen Concentration. 

To allow for a fair comparison between different flight tests, it is critical to express the oxygen 

concentration of the tank as a whole, even though the concentrations of the individual bays often 

vary. To achieve this, a weighted average by volume was calculated, given the oxygen 

concentration distribution at a given time. This average weighed the oxygen concentration of 

each bay with the total tank volume percentage of each bay. 

 (2) 0.10[O 4] 0.13[O 5] 0.13[O 6]

[O ] 31[O 1] 0.23[O 2] 0.10[O 3]

2 2 2

2 2 2 2

Bay Bay Bay

Average . Bay Bay Bay

+ + +

= + +

4.1.3 Inerting Model. 

A simple analytical model was developed to calculate the average tank ullage oxygen 

concentration, given a specific tank volume and system performance schedule, to compare to the 

measured tank average oxygen concentration data. The model calculates the mass of oxygen in

the tank at the start of the mission, given a starting tank oxygen concentration, and tracks the 

mass of oxygen in and out of the tank, given the changing system performance and flight 

conditions including the appropriate net oxygen entering the tank as a result of descent. The 

following equation governs the model process. 

( ) ( 1) ( 1) ( ) ( 1) ( ) 21 O2 O2 m t m t m IGOF m UGOF t *V UGOF t *V * . = − + ∗ − ∗ − − ∆ρ Tank ∗ − + ∆ρ Tank & & (3) 

24
In this equation, UGOF(t-1) is the fraction of oxygen in the ullage gas. It is calculated by 

dividing the mass of oxygen in the tank at t-1 by the mass of gas in the tank ullage or: 

( 1) ( 1) ( 1) (4) O2UGOF t − = m t − / mTank t −

Where: ( ) = Mass of oxygen in tank at time t O2

m t

 m& = Mass flow rate of inerting gas (in terms of t) 

 IGOF = Fraction of oxygen in inerting gas 

 ∆ρ = Change in ullage density due to altitude change 

 VTank = Volume of tank ullage 

 mTank = Mass of gas in tank 

 mair = Mass of air entering tank 

A more detailed description of the model process is given in reference 5. 

4.2 FLAMMABILITY CALCULATIONS.

The primary calculations performed to analyze the flammability test data were simplified fuel air 

ratio (FAR) calculations using an FAA-developed model. Additionally, a more complex ullage

flammability model was employed. 

4.2.1 Estimation of the Vapor Generation in the CWT.

The vapor generation model used by the FAA has been developed by Professor C. E. 

Polymeropoulos of Rutgers University, and is still in its validation phase. The model employs 

free convection heat and mass transfer correlations in a fully mixed tank. It requires 

experimental data on the liquid fuel temperature, the tank walls, the ambient temperature and

pressure, and generates, numerically over time, the total mass of vapor generated and the vapor

masses of the component species used to characterize the liquid fuel. The user is able to input 

pressure, liquid fuel, and wall temperatures as functions of time, and can select from the 

compositions of several different flashpoint fuels. Part of the model can also calculate the 

equilibrium vapor concentrations at different pressures and temperatures. More detailed 

information about this model can be found in reference 8. 

4.2.2 Fuel Air Ratio Calculator.

The FAR calculator, developed by the FAA’s Chief Scientific and Technical Advisor for fuel 

system design, is by comparison a much simpler model, but it still provides a great amount of 

useful data. This model predicts fuel vapor pressure, and therefore, FAR for a wide range of 

fuels over a wide range of altitudes, temperatures, and mass loadings. This model, however,

only calculates equilibrium values occurring under isothermal conditions and, therefore, is a 

conservative estimate of the FAR. The model also offers several other features, such as being 

able to enter ASTM D 2887 distillation curve and flashpoint data to generate a customized fuel 

for use in the model and predicting such things as the flashpoint of a given fuel, molecular 

weight of the vapor, and the temperature needed to arrive at a stoichiometric mixture. The model

25
and further associated information can be found online at www.fire.tc.faa.gov/systems/

fueltank/downloads.stm. 

5. DISCUSSION OF RESULTS. 

The results of the flight tests are presented for two areas of analysis: fuel tank inerting and fuel

tank flammability. 

5.1 FUEL TANK INERTING. 

The primary focus of the fuel tank inerting data is the changing system performance, given 

different flow modes and orifice settings. Additionally, the effect of these flow conditions on 

OBIGGS bleed air consumption was examined. Also, the ability of the system to inert the given

aircraft CWT was examined for both the average fuel tank oxygen concentration and the 

distribution of oxygen within the compartmentalized tank. 

5.1.1 Inerting System Performance.

The primary factors affecting the inerting system flow and purity are the pressure on the system 

and the altitude (pressure) at the OEA vent. Figure 18 illustrates these two parameters for the 

OBIGGS NEA flow and purity for a test using the basic dual-flow methodology. As expected, 

increases in pressure on the ASM resulted in greater NEA flow and a lower NEA oxygen 

concentration (more pure), given a fixed orifice. The NEA oxygen concentration will decrease 

with increased flow when the NEA deposit orifice is fixed. This is due to the fact that an

increase in flow across a fixed orifice will result in a greater pressure drop and an increase in

pressure drop across this system flow orifice will result in a decrease in oxygen concentration. 

0

5

10

15

20

25

0 20 40 60 80 100 120 140 160

Time (mins)

Oxygen 

Concentration (% vol)

0

10

20

30

40

50

60

70

Alt (1000 ft) \ Pressure (psia) \ Flow (scf

m)

 NEA [O2] NEA Flow

 ASM Pressure Altitude

Test 1 Data - OBIGGS Performance

FIGURE 18. SYSTEM PERFORMANCE DATA FOR A TYPICAL FLIGHT TEST

26
Figure 19 illustrates the ASM pressure for the taxi/takeoff portion of three different tests with

three different ascent rates. This graph illustrates the very similar pressure profiles the system 

was exposed to at the inception of each test, particularly the period just after takeoff for 

approximately 4 minutes. Figures 20 and 21 illustrate the NEA purity and flow, respectively, for

the same three tests during the same time period. Each test had an adjusted time to allow for

comparison of the parameters. Figures 20 and 21 illustrate the somewhat wider range of

performance for this same time period. 

0

10

20

30

40

50

60

70

15 20 25 30 35 40 45

Adjusted Time (mins)

Pressure (psia)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Test 1 ASM Pressure

Test 2 ASM Pressure

Test 4 ASM Pressure

Test 1 Altitude

Test 2 Altitude

Test 4 Altitude

ASM Pressure Comparison

FIGURE 19. COMPARISON OF ASM PRESSURE DURING THE INCEPTION OF THREE 

DIFFERENT FLIGHT TESTS WITH ALTITUDE 

Further examination of the system input conditions illustrates that the system input temperature 

was consistent for each test, but each had a significantly different output temperature, indicating

a different level of heat soaking in the ASMs. For each test, the system was run prior to the test,

but for different periods of time, each test having the system off prior to taxi (system on for test) 

for a different amount time. Also, the ACMs were run in different ways prior to each test,

creating different temperature conditions in the pack bay area where the system was housed. All 

these factors had an effect on how heat soaked (steady temperature operation) the ASMs are, 

which created these unique performance conditions for the initial part of each system operation. 

This is evident in figure 22, which illustrates a similar ASM temperature input range for each of 

the tests during this time period (21 to 25 minutes) but very different NEA temperatures (ASM 

output). ASM temperature changes the permeability characteristics of the ASM, which causes

these marked changes in system performance observed in figures 20 and 21. Presumably, these 

variations would be greater if the ambient temperatures were colder. Ambient temperatures

ranged from 80o to 95oF with varying amounts of sun and wind for these three tests. 

27
0

2

4

6

8

10

12

14

16

18

20

15 20 25 30 35 40 45

Adjusted Time (mins)

O

xygen C

oncentration (v

ol %)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Test 1 NEA [O2]

Test 2 NEA [O2]

Test 4 NEA [O2]

Test 1 Altitude

Test 2 Altitude

Test 4 Altitude

NEA [O2] Comparison

FIGURE 20. COMPARISON OF NEA PURITY DURING THE INCEPTION OF THREE 

DIFFERENT FLIGHT TESTS WITH ALTITUDE 

0

5

10

15

20

25

30

35

40

15 20 25 30 35 40 45

Adjusted Time (mins)

Flo

w (scfm)

0

5

10

15

20

25

30

35

40

Altitu

d

e (1000 ft)

Test 1 NEA Flow

Test 2 NEA Flow

Test 4 NEA Flow

Test 1 Altitude

Test 2 Altitude

Test 4 Altitude

NEA Flow Comparison

FIGURE 21. COMPARISON OF NEA FLOW DURING THE INCEPTION OF THREE 

DIFFERENT FLIGHT TESTS WITH ALTITUDE 

28
Figure 23 illustrates the ASM inlet pressure for the descent/landing portion of three different 

tests with three similar descent profiles. Each test used a different valve and methodology for

the variable/high-flow portion of the flight. As previously discussed, test 1 had a fixed high-flow 

orifice, while tests 2 and 3 varied the high-flow orifice to maximize flow at the top of descent. 

This graph illustrates the very similar pressure profiles the system was exposed to at the 

beginning of each descent, marked by periods of high variation in pressure during the altitude

hold and completion of descent. All three tests exhibited the same range of pressures with 

similar trends. 

0

50

100

150

200

250

15 20 25 30 35 40 45

Adjusted Time (mins)

Tem

perature (deg F)

Test 1 ASM Temp Test 1 NEA Temp

Test 2 ASM Temp Test 2 NEA Temp

Test 4 ASM Temp Test 4 NEA Temp

System Temperature Comparison

FIGURE 22. COMPARISON OF ASM INPUT AND OUTPUT CONDITIONS DURING THE 

INCEPTION OF THREE DIFFERENT FLIGHT TESTS 

Figures 24 and 25 illustrate the NEA purity and flow, respectively, for the same three tests 

illustrated in figure 23. Each test had an adjusted time to allow for comparison of the 

parameters. Figures 24 and 25 show the expected result that tests 2 and 3, with potentially wider 

high-flow orifices, delivered a resulting higher volume of NEA with a resulting higher oxygen 

concentration (less pure), particularly at the bottom of descent. The result of this system 

performance change should be less ambient air flowing into the CWT through the vent system, 

and thus, a net lower oxygen concentration in the tank after the completion of descent. Note the 

erratic nature of the ASM pressure and flow at the end of descent, making it difficult to judge 

what tests had a greater net flow into the CWT. This is due to the very dynamic nature of the 

engine throttle settings during the final stages of flight. 

29
0

10

20

30

40

50

60

105 110 115 120 125 130 135 140 145 150 155

Adjusted Time (mins)

Pressure (psia)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Test 1 ASM Pres

Test 2 ASM Pres

Test 3 ASM Pres

Test 1 Altitude

Test 2 Altitude

Test 3 Altitude

ASM Pressure Comparison

FIGURE 23. COMPARISON OF ASM PRESSURE DURING THE DESCENT PHASE OF 

THREE DIFFERENT FLIGHT TESTS WITH ALTITUDE 

0

2

4

6

8

10

12

14

16

18

20

105 110 115 120 125 130 135 140 145 150 155

Adjusted Time (mins)

O

x

y

gen Concentr

ation (% vol)

0

5

10

15

20

25

30

35

40

Altitu

d

e (1000 ft)

Test 1 NEA [O2]

Test 2 NEA [O2]

Test 3 NEA [O2]

Test 1 Altitude

Test 2 Altitude

Test 3 Altitude

NEA [O2] Comparison

FIGURE 24. COMPARISON OF NEA PURITY DURING THE DESCENT PHASE OF 

THREE DIFFERENT FLIGHT TESTS WITH ALTITUDE 

30
0

10

20

30

40

50

60

70

80

105 110 115 120 125 130 135 140 145 150 155

Adjusted Time (mins)

Altitu

d

e (kft) / Fl

o

w (scfm)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Test 1 NEA Flow

Test 2 NEA Flow

Test 3 NEA Flow

Test 1 Altitude

Test 2 Altitude

Test 3 Altitude

NEA Flow Comparison

FIGURE 25. COMPARISON OF NEA FLOW DURING THE DESCENT OF THREE 

DIFFERENT FLIGHT TESTS WITH ALTITUDE 

Bleed air flow consumption for the taxi/takeoff and descent/landing phases of the abovementioned tests was also examined. The taxi/takeoff bleed air consumption (figure 26) shows a

marked variation between the three tests, similar to the comparison of NEA flow. This is 

consistent with the analysis that temperature variations in the ASMs caused these differences.

Any difference in ASM temperature will manifest itself in terms of changes in permeability, 

which will affect the NEA flow (recovery) by changing the bleed air consumption 

(permeability). The descent/landing bleed air consumption (figure 27) shows no significant

deviation between the three tests, unlike the comparison of NEA flow. This is consistent with 

the ASM permeability characteristics that generally give constant ASM input flow, given a fixed 

ASM pressure and permeate pressure, assuming a heat-soaked bundle at a constant temperature. 

With each test operating with the same pack settings at the same altitude conditions for 

approximately the same time, somewhat constant temperature conditions would be expected. 

The NEA flow is a function of orifice pressure drop, as dictated by orifice size. The differences

shown in NEA flow in figure 25, during the second half of the first part of descent, are due to 

changes in the high-flow orifice size from test to test, or by mechanically changing the effective 

orifice diameter during a test to increase flow and decrease purity. These tend to have little or no 

effect on the bleed air consumed by the OBIGGS during the observed high-flow conditions. 

31
0

20

40

60

80

100

120

140

160

180

15 20 25 30 35 40 45

Adjusted Time (mins)

Flow (scfm)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Test 1 Bleedair Flow

Test 2 Bleedair Flow

Test 4 Bleedair Flow

Test 1 Altitude

Test 2 Altitude

Test 4 Altitude

Bleedair Flow Comparison

FIGURE 26. COMPARISON OF BLEED AIR CONSUMPTION DURING THE INCEPTION 

OF THREE DIFFERENT FLIGHT TESTS WITH ALTITUDE 

0

50

100

150

200

250

300

105 110 115 120 125 130 135 140 145 150 155

Adjusted Time (mins)

Flow (scfm)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Test 1 Bleedair Flow

Test 2 Bleedair Flow

Test 3 Bleedair Flow

Test 1 Altitude

Test 2 Altitude

Test 3 Altitude

Bleedair Flow Comparison

FIGURE 27. COMPARISON OF BLEED AIR FLOW DURING THE DESCENT OF THREE 

DIFFERENT FLIGHT TESTS WITH ALTITUDE 

32
A correlation of system NEA flow and pressure was made for selected portions of the ascent and 

descent to characterize the system performance. It is important to note that these relationships 

are only valid for a single orifice setting and a narrow range of altitudes. Each specified altitude

relates to a wide range of data around that altitude to make generalizations about system flow 

versus ASM pressure. Figure 28 shows how sensitive the NEA flow is to ASM pressure, even in

the low-flow mode. The excellent correlation was due to the fact that the low-flow orifice 

settings remained constant during the entire flight test. The deviations in slope between the data 

from test 1 and the test 2 and 4 are probably due to the difference in system warmup times

causing the deviations in system performance. 

0

10

20

30

40

50

60

70

5 10 15 20 25

Flow (scfm)

Pressure (psia)

30

 Test 1 - Approx 15k Feet

 Test 2 - Approx 15k Feet

 Test 4 - Approx 15k feet

 Test 1 - Sea Level

 Test 2 - Sea Level

 Test 4 - Sea Level

 Test 1- Approx 25k Feet

 Test 2 - Approx 25k Feet

 Test 4 - Approx 25k Feet

ASM Pressure / NEA Flow Correlation

FIGURE 28. CORRELATION OF NEA FLOW WITH ASM PRESSURE FOR THE SYSTEM

LOW-FLOW MODE FOR THREE DIFFERENT FLIGHT TESTS 

Figure 29 gives the correlation between ASM pressure and NEA flow using the variable/highflow mode. Although the data represents a wide range in orifice settings, better correlation is 

expected. The poor correlation in the graph is most likely due to the difficulty in obtaining 

consistent ASM pressure and orifice setting during the descent. The advantage of the variableflow orifice used by the system operator for tests 2 through 5 was difficult to gage because of 

problems obtaining stable observable parameters (NEA purity, NEA flow, ASM pressure) with 

which to schedule the orifice setting. These problems were exacerbated because the operator did

not have the ability to gage the existing orifice setting during the descent, but rather, changed the 

orifice setting continually to obtain the flow and purity desired. 

33
0

10

20

30

40

50

60

5 10 15 20 25 30 35

Flow (scfm)

Pressure (psia)

40

 Test 1 - Sea Level

 Test 2 - Sea Level

 Test 3 - Sea Level

 Test 1 - Approx 15k Feet

 Test 2 - Approx 15k Feet

 Test 3 - Approx 15k feet

 Test 1 - Approx 25k Feet

 Test 2 - Approx 25k Feet

 Test 3 - Approx 25k Feet

ASM Pressure / Flow Correlation

FIGURE 29. CORRELATION OF NEA FLOW WITH ASM PRESSURE FOR THE SYSTEM

VARIABLE/HIGH-FLOW MODE FOR THREE DIFFERENT FLIGHT TESTS 

5.1.2 Ullage Oxygen Concentration Data.

The result of the system performance on the oxygen concentration in the ullage of the CWT was 

analyzed by examining the oxygen concentration in each of the six bays in the tank. Figure 30 

shows that there is considerable difference in the highest and lowest oxygen concentration in the 

ullage, particularly during descent. Bay 1 has a consistently higher oxygen concentration than 

the rest of the bays, primarily because it is the farthest from the bay where the NEA is deposited

(bay 6), and it is where most of the air is deposited from the vent system during descent. 

5.1.2.1 Average/Worst Bay Oxygen Concentration.

The primary focus of the fuel tank inerting analysis was to determine the ability of the system to 

maintain an average ullage oxygen concentration less than 12 percent oxygen in the CWT. 

Figure 31 gives the average ullage oxygen concentration measured in the CWT, as calculated 

from the data in figure 30, with the altitude flight profile for a reference. This profile was typical 

of the tests performed that had the tank oxygen concentration starting between 9 and 12 percent 

oxygen, which quickly decreased to about 3 percent after approximately 1 hour of cruise. At the

time of descent, the oxygen concentration in the tank spikes, due to air entering the tank, and 

each hold at altitude allows for the tank oxygen concentration to decrease again. The average 

oxygen concentration in the tank settles out briefly after touchdown. It is important to note how

quickly the oxygen concentration in the tank drops after takeoff. Although the normalized 

volume flow of NEA from the system in low-flow mode only changes a relatively small amount

from taxi and takeoff to cruise, the impact of the actual volume flow at cruise altitude on the 

CWT oxygen concentration was much greater than at sea level when considering the mass of gas 

in the tank (to be ullage washed) is less than 1/3 the mass at sea level. 

34
0

5

10

15

20

25

0 20 40 60 80 100 120 140 160

Time (mins)

Oxygen Concentrati

on (% vol)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Bay 1 Bay 2

Bay 3 Bay 4

Bay 5 Bay 6

Altitude

Test 3 Data - System Variation 2

FIGURE 30. OXYGEN CONCENTRATION DATA FROM ALL SIX BAYS WITH TEST 

ALTITUDE FOR A TYPICAL FLIGHT TEST 

0

5

10

15

20

25

0 20 40 60 80 100 120 140 160

Time (mins)

Oxygen Concentrati

on (% v

ol)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

Average Ullage

Altitude

Test 3 Data - System Variation 2

FIGURE 31. AVERAGE ULLAGE OXYGEN CONCENTRATION DATA CALCULATED

FROM SIX BAY MEASUREMENTS WITH TEST ALTITUDE FOR A 

TYPICAL FLIGHT TEST 

35
Figure 32 compares four descent profiles (altitude) with the measured average ullage oxygen 

concentration. Each test has a different flight profile with similar features. As evident from the 

graph, the total change in altitude (pressure) and the time to descend are the key factors when 

considering the increase in ullage oxygen concentration during descent. The same can be said of 

the worst bay oxygen concentration, described previously as bay 1. Figure 33 illustrates the bay 

1 oxygen concentration for the same four descent profiles and illustrates the very same trends 

observed in figure 32 with greater magnitude. Although the greatest observed ullage average 

peak was 13.5 percent for a descent from 35,000 feet in 41 minutes, the corresponding bay 1 

peak was approximately 19.5 percent oxygen by volume. However, on a 64-minute descent

from 33,000 feet, the average ullage oxygen concentration peak was estimated to be 

approximately 10 percent, while the corresponding bay 1 peak was estimated to be less than 14 

percent by volume. The peaks for test 5 were estimated due to a loss of data during the final

stages of the flight. These estimates were made by calculating the peak oxygen concentration 

using a simple ullage model and observing the peaks of a similar flight test. This illustrates that

an effective way to decrease the peak oxygen concentration in the worst observed bay is to 

decrease overall average ullage oxygen concentration at the duration of the flight. This is 

illustrated in figure 34, which gives the average ullage oxygen concentration correlated with the 

worst bay oxygen concentration for the four tests presented in figures 32 and 33. 

Estimated

Test 5 Peak

0

2

4

6

8

10

12

14

16

0 10 20 30 40 50 60 70 80 90

Adjusted Time (mins)

O

xygen Concentration (% v

ol)

0

5

10

15

20

25

30

35

40

Altitude (1000 ft)

 Test 1 Average O2

 Test 3 Average O2

 Test 4 Average O2

 Test 5 Average O2

 Test 1 Altitude

 Test 3 Altitude

 Test 4 Altitude

 Test 5 Altitude

Comparison of Descent Profiles

with Average Ullage [O2]

FIGURE 32. COMPARSION OF AVERAGE ULLAGE OXYGEN CONCENTRATION 

WITH FOUR DIFFERENT TEST DESCENT PROFILES 

36
Estimated

Test 5 Peak

0

5

10

15

20

25

0 10 20 30 40 50 60 70 80 90

Adjusted Time (mins)

Oxygen Concentration (% vol)

0

5

10

15

20

25

30

35

40

Altitu

d

e (1000 ft)

 Test 1 Worst Bay O2 Test 3 Worst Bay O2

 Test 4 Worst Bay O2 Test 5 Worst Bay O2

 Test 1 Altitude Test 3 Altitude

 Test 4 Altitude Test 5 Altitude

Comparison of Descent Profiles

with Worst Bay [O2]

FIGURE 33. COMPARSION OF THE WORST BAY OXYGEN CONCENTRATION WITH 

FOUR DIFFERENT TEST DESCENT PROFILES 

0

2

4

6

8

10

12

14

16

18

20

0 2 4 6 8 10 12 14 16

Average Ullage Peak [O2] (% vol)

Worst Bay Peak [O2] (% vol)

Worst Bay / Average Ullage [O2] Correlation

4 Different Test Descents

FIGURE 34. CORRELATION OF THE PEAK WORST BAY OXYGEN CONCENTRATION 

WITH THE PEAK AVERAGE ULLAGE OXYGEN CONCENTRATION FOR FOUR 

DIFFERENT TEST DESCENT PROFILES 

37
5.1.2.2 Inerting Model Comparison.

To gage the ability of a simple ullage inerting model to predict average oxygen concentration, 

the model discussed in section 4.1.3 was applied to the data. The model uses the measured 

system performance and flight data to generate the ullage oxygen concentration. The model 

assumes a single, well-mixed bay in the tank vented to atmospheric pressure. Figure 35 shows 

the results of the model compared to the average ullage oxygen concentration calculated from the

individual bay measurements. A comparison of the data shows good agreement with the data 

trends, with some disagreement between the peak oxygen concentrations. The maximum and 

minimum values agree within 1 oxygen percent, with some discrepancy between the time at

which the peaks occur. Any deviation in peak times can be explained by the long delay in the 

OBOAS sample system, which takes as much as 2-3 minutes to completely respond to rapid 

changes in the ullage oxygen concentration. The differences in peak values can be explained by 

the efficiency of the inerting process when comparing a single deposit inerting a six-bay tank 

versus a single-bay rectangular tank. Previous studies of inerting compartmentalized fuel tank 

models on the ground exhibited an increase in inerting efficiency when using single- deposit 

methods [9]. This increased efficiency would manifest itself as a lower oxygen concentration 

achieved for the same gas deposited, similar to what is observed in figure 35. 

0

5

10

15

20

25

0 20 40 60 80 100 120 140 160 180

Time (mins)

Oxygen Con

c

e

ntrati

on (v

ol %)

0

2

4

6

8

10

12

14

16

Pressure (psia)

 Inerting Model

 NASA Data

 Ambient Pressure

NASA Data - Dual Mode

System

FIGURE 35. AVERAGE ULLAGE OXYGEN CONCENTRATION DATA FOR A TYPICAL 

FLIGHT TEST COMPARED TO A SIMPLE ULLAGE INERTING MODEL 

38
5.1.2.3 Inert Gas Distribution.

The most challenging part of the distribution of inert gas in the CWT was during descent. Figure 

36 shows the measured oxygen concentration in six bays of the CWT while the inerting system 

was operating in the variable-flow mode. This mode made it possible to create even higher 

flows at the top of the descent than the ordinary high-flow mode. Although most tests varied in 

starting altitude and descent time, they all exhibited similar trends with comparable distribution 

of the inert gas at the end of the test, with respect to the different bay oxygen concentrations. 

0

5

10

15

20

25

105 110 115 120 125 130 135 140 145 150 155

Time (mins)

Oxygen Concentrati

o

n (

% vol)

0

5

10

15

20

25

30

35

40

Altitu

d

e (1000 ft)

Bay 1 Bay 2 Bay 3

Bay 4 Bay 5 Bay 6

Altitude

Test 3 Data - System Variation 2

FIGURE 36. OXYGEN CONCENTRATION DATA FOR THE SIX BAYS OF THE CWT 

DURING THE DESCENT OF A TYPICAL FLIGHT TEST

Three different system operation methodologies were used for three different test descents, 

numbered test 1, 2, and 3. Test 1 used the traditional dual-flow mode, while tests 2 and 3 used 

two different scheduling methods for the variable-flow mode. Test 3 used the variable-flow

mode to maximize flow, while test 2 used the variable-flow mode to target a particular oxygen 

concentration, which tended to give more flow than test 1. Figure 37 compares the distribution

of inert gas at the end of each test, 5-10 minutes after touchdown. Each test started descent with 

approximately 2.5 percent average ullage oxygen concentration. A subtle trend was the variableflow mode seemed to improve distribution, but it was difficult to tell as all three tests did not use 

the same descent time. Qualitatively, it appears that if test 1 had a 41-minute descent as in the 

case of tests 2 and 3, test 1 would have a greater average ullage oxygen concentration and a 

greater bay 1 (worst bay) oxygen concentration. This would imply that the more flow on 

descent, the better the distribution of inert gas at the completion of descent, but apparently this

had little or no effect on the average ullage oxygen concentration after touchdown. 

39
0

2

4

6

8

10

12

14

O

x

ygen Concentration (

% vol)

 Bay 1

 Bay 2

 Bay 3

 Bay 4

 Bay 5

 Bay 6

 Average

Test 1 Test 2 Test 3 46 Minute Descent 41 Minute Descent 41 Minute Descent

FIGURE 37. VARIATION IN THE OXYGEN CONCENTRATIONS IN THE SIX BAYS OF 

THE CWT DURING THREE FLIGHT TESTS 

5.1.2.4 Inert Gas Diffusion.

To determine the potential flammability exposure of an aircraft landing with one or more tank 

bays above 12 percent oxygen, but with the average tank ullage oxygen concentration being less

than 12 percent, a 3-hour ground turnaround test was performed in a warm dry climate 

(figure 38). The aircraft was parked after landing and the instrumentation was turned off briefly 

for maintenance purposes. One hour later, the instrumentation was turned on again and the 

oxygen concentration of the ullage was monitored. It took approximately 1 hour for the worst

bay (bay 1) to go from 13 percent to 12 percent, with an average ullage oxygen concentration of

11 percent oxygen. During the ground turnaround, the ACMs were operated continuously to 

keep the cabin air conditioned, which heated the CWT. The average ullage oxygen 

concentration increased from approximately 10 percent at time zero, just after landing, to 

approximately 11 percent at time 1 hour. This is most likely due to slight vertical gradients in

the individual bay oxygen concentrations from the mixing of different temperatures of ullage 

gases, NEA, and ambient air. Since the ullage gas samples are taken entirely at the top of the 

tank, any vertical nonhomogeneities in a bay will result in a biased reading. Similar behavior 

was observed on a 747SP ground test article used by the FAA for ground-based inerting and 

OBIGGS development and testing [10]. In this case, large quantities of NEA were deposited in a 

quiescent tank over a relatively short period of time while oxygen samples were taken 

continuously. After the NEA was stopped and a brief settling period ensued, the ACMs were 

turned on, and the average tank ullage oxygen concentration would rise and then stabilize over a 

period of several minutes. Any nonhomogeneities occurring during the flight test, did so during 

ACM operation, but were apparently more subtle in nature. 

40
0

2

4

6

8

10

12

14

0 0.5 1 1.5 2 2.5 3 3

Time (hours)

Oxygen Concentration (

% vol)

.5

 Bay 1 Bay 2 Bay 3

 Bay 4 Bay 5 Bay 6

 Average

747 CWT Bay-to-Bay Oxygen Concentration 

Distribution for a 3-Hour Turn-Around

FIGURE 38. COMPARSION OF THE ULLAGE OXYGEN CONCENTRATION FOR ALL 

SIX BAYS DURING A 3-HOUR TURNAROUND 

To examine the effects of air entering the ullage during overnight periods, the ullage oxygen 

concentration and distribution was recorded at the completion of a test, and then again the next

morning after the instrumentation was warmed up and calibrated. The CWT was not crossvented to both wing vent tanks, as is typical of the 747 model. Cross venting has been associated 

with air ventilating the inert ullage of Boeing type aircraft [4], and a modification was performed 

on the vent system to eliminate this venting process (section 2.1.1.2). Figure 39 shows the 

distribution of oxygen concentration of the six bays of the tank after touchdown one afternoon 

and then again early the next morning. As expected, the distribution of bay oxygen 

concentrations came completely to equilibrium. During the 12-hour overnight period, the 

average ullage oxygen concentration rose slightly over 1 percent, even though fuel was 

transferred during the instrumentation warmup period. Assuming the time zero average oxygen

concentration is an unbiased reading and the fuel transfer deposited no air in the tank, this still 

represents a small increase in ullage oxygen concentration. This is consistent with the theory

that thermal cycling of the tank is the source of air penetration because this particular night was 

almost as warm as the previous day. It is just as likely that virtually no air entered the tank

during the overnight period, and the increase in oxygen concentration is strictly due to the final 

equilibration of the inert gas in the tank, as observed in the first hour of the 3-hour turnaround 

test and/or the transfer of fuel from tanks 2 and 3. 

41
0

2

4

6

8

10

12

14

0 2 4 6 8 10 12 14 1

Time (hours)

Oxy

gen Concentration (% vol)

6

 Bay 1 Bay 2 Bay 3

 Bay 4 Bay 5 Bay 6

 Average

747 CWT Bay-to-Bay Oxygen Concentration 

Distribution for a 12-Hour Overnight Sit with

Fuel Transfer

25% Fuel Load Transferred

to CWT from Tanks 2 & 3

FIGURE 39. COMPARSION OF OXYGEN CONCENTRATION FOR ALL SIX BAYS 

DURING A 24-HOUR OVERNIGHT PERIOD WITH A FUEL TRANSFER 

5.2 FUEL TANK FLAMMABILITY.

The primary focus of the fuel tank flammability measurements was to generate real-time, inflight THC data for comparison to predicted results from the FAA’s fuel tank flammability

models. The results of these measurements, including environmental and flight effects are first 

discussed in the following sections followed by a comparison of the data to the calculated results. 

The flashpoint and ASTM D 2887 distillation data on the fuels for each test are shown in 

appendix I. 

5.2.1 General Flammability Trends. 

Though the data from each flight test is quite different, there are certain trends in the data that are

seen throughout. As an example of these trends, figure 40 shows the THC and ambient pressure 

data for both the CWT and the wing tank from flight test 5. Initially, during ground operations 

with the ACMs on, the CWT THC reading rises slowly, but steadily, prior to takeoff. During 

takeoff, the effects of the change in vapor pressure are seen, as the hydrocarbons evolve faster, 

overcoming the corresponding condensation effects due to reduced temperatures. This effect is 

reversed at the approximate time of the start of level flight, as the condensation effects now

begin to drive the THC readings. Finally, during descent, incoming air causes the THC to drop

at a slightly increased rate. Similar trends are seen in the wing tank flammability data; however,

the condensation effects in this tank are much more substantial as its ullage tends to follow the 

ambient temperature more closely. As a result, there is little if any increase in the wing tank

flammability measurements during the climb phase of flight.

42
0.0

0.5

1.0

1.5

2.0

2.5

3.0

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00 4:30 5:00 5:30 6:00 6:30

Test Time

THC (% Propane)

0

3

6

9

12

15

18

Pressure (psia)

CWT (% PROPANE)

WT2 (% PROPANE)

Ambient Pressure

FIGURE 40. FLIGHT TEST 5 THC AND AMBIENT PRESSURE MEASUREMENTS 

5.2.2 Temperature Effects.

Temperature effects on the THC readings throughout testing were quite noticeable in both tanks. 

Figures 41 and 42 show the temperature and THC data from flight test 5 for the CWT and wing 

tank, respectively. The approximate times of takeoff and start of descent are labeled on each for 

reference. In each figure, it is clear that the THC readings in each tank closely follow the trends

of temperature. The only exception to this is in the CWT during the climb phase of flight, as the 

vapor pressure effects overcome the effects of condensation. During ground operations prior to 

takeoff, as fuel, surface, and ullage temperatures slowly increase, both from operation of the 

ACMs and ambient heating, so does the THC readings in each tank. In the cruise phase of flight, 

as temperatures drop due to the cold ambient temperature, the THC readings closely follow the 

trend of all fuel tank temperatures as condensation causes the THC levels to decline. Likewise, 

as temperatures in the wing tank begin to increase once again during descent and landing, so too 

does the wing tank flammability. 

43
0

20

40

60

80

100

120

140

160

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00 4:30 5:00 5:30 6:00

Test Time

Temperature (F)

0

0.3

0.6

0.9

1.2

1.5

1.8

2.1

2.4

THC (% Propane)

Avg Ullage (°F)

Avg Floor (°F)

Avg Wall (°F)

Sump Fuel Bay (°F)

Avg Ceiling (°F)

CWT (% Propane)

Approximate 

Time of Take-off

Approximate Time

of Start of Descent

FIGURE 41. FLIGHT TEST 5 CWT THC AND TEMPERATURE MEASUREMENTS 

-20

0

20

40

60

80

100

120

140

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00 4:30 5:00 5:30 6:00

Test Time

Temperature (F)

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

THC (

% Propane)

#2WT FWD Spar Surf (°F)

#2WT Inboard Fuel (°F)

#2WT Rear Spar Surf (°F)

#2WT Ullage (°F)

#2WT O/B Wall Surf (°F)

#2WT Top Surf (°F)

WT2 (% Propane)

Approximate Time

of Take-off

Approximate Time

of Start of Descent

FIGURE 42. FLIGHT TEST 5 WING TANK THC AND TEMPERATURE 

MEASUREMENTS 

44
To further exhibit the close relationship between tank flammability and various tank

temperatures, figure 43 shows the CWT temperature and THC data from flight test 4. Because 

this test was performed with a 25% fuel load in the CWT, the changes in temperatures in flight 

are relatively small, with the exception of the tank wall temperatures. As the tank temperatures

remain relatively stable, so do the THC readings in the tank, although it is evident that the drop 

in the tank wall temperatures still has an effect on the THC, causing a slow decrease in the 

flammability readings throughout the cruise phase of flight.

0

30

60

90

120

150

180

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00 4:30 5:00

Test Time

Temperature (F)

0

0.5

1

1.5

2

2.5

3

THC (% Propane)

Avg Ullage (°F)

Avg Floor (°F)

Avg Wall (°F)

Sump Fuel Bay 1 (°F)

Avg Ceiling (°F)

CWT (% Propane)

Approximate 

Time of Take-off

Approximate Time

of Start of Descent

FIGURE 43. FLIGHT TEST 4 CWT THC AND TEMPERATURE MEASUREMENTS 

5.2.3 Cross-Venting Effects.

The effects of cross venting (see section 2.1.1.2) on CWT flammability are shown in figures 44 

and 45. These figures depict CWT THC and ambient pressure readings in tests 0 and 6, 

respectively. Both flights consisted of a similar flight profile without the OBIGGS running and 

showed similar trends in CWT temperature readings, as shown in figures 46 and 47. 

45
0

0.5

1

1.5

2

2.5

3

3.5

4

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00

Test Time

THC (

% Propane)

0

2

4

6

8

10

12

14

16

Pressure (psia)

CWT (% Propane)

Ambient Pressure (psia)

FIGURE 44. FLIGHT TEST 0 CWT THC AND AMBIENT PRESSURE MEASUREMENTS 

0

0.5

1

1.5

2

2.5

3

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00 4:30 5:00 5:30 6:00 6:30 7:00 7:30

Test Time

THC (

% Propane)

0

3

6

9

12

15

18

Pressure (psia)

CWT (% PROPANE)

Ambient Pressure (psia)

FIGURE 45. FLIGHT TEST 6 CWT THC AND AMBIENT PRESSURE MEASUREMENTS 

46
0

20

40

60

80

100

120

140

160

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00

Test Time

Temperature (F)

Avg Ullage (°F)

Avg Floor (°F)

Avg Wall (°F)

Avg Ceiling (°F)

Sump Fuel Bay 1 (°F)

Approximate 

Time of Take-off Approximate Time

of Start of Descent

FIGURE 46. FLIGHT TEST 0 CWT TEMPERATURE MEASUREMENTS 

0

20

40

60

80

100

120

140

160

0:00 0:30 1:00 1:30 2:00 2:30 3:00 3:30 4:00 4:30 5:00 5:30 6:00 6:30 7:00 7:30

Test Time

Temperature (F)

Avg Ullage (°F)

Avg Floor (°F)

Avg Wall (°F)

Sump Fuel Bay 1 (°F)

Avg Ceiling (°F)

Approximate 

Time of Take-off Approximate Time

of Start of Descent

FIGURE 47. FLIGHT TEST 6 CWT TEMPERATURE MEASUREMENTS 

47
48

Flight test 0 was conducted with half of the venting system blocked, eliminating all cross 

venting, whereas the venting system was in the original configuration through flight test 6, thus 

allowing cross flow through the CWT. Data from flight test 0 was somewhat scattered due to 

various issues during the flight, which caused the system to be diverted from fuel tank sampling 

at several points through the test. Using the data and test notes taken during the flight, a trend 

line of the data during the cruise phase of flight has been drawn, depicting the trend of the 

available data. Pressure data from test 6 was unavailable through much of the flight because the 

ambient pressure transducer failed shortly after takeoff. It was noted however, that the cruise 

altitude was approximately 31,000 ft (4.17 psia) for nearly the entire flight. 

From figures 44 and 45, the same trends discussed in section 5.2.1 are shown; however, during 

cruise, the CWT flammability readings drop much more rapidly when cross flow is present, 

despite the similarities in flight profile and tank temperature data. It is also interesting to note 

that these two tests, the only two conducted without the OBIGGS running, exhibited the highest 

peak THC readings. Presumably, when the OBIGGS is operating, the incoming NEA displaces a 

sufficiently high amount of ullage vapors to produce a net lower flammability in the tank. 

Therefore, the OBIGGS can potentially protect against ignition not only by reducing the fuel 

tank ullage oxygen concentration, but also by reducing the resulting fuel/air mass ratio in the 

ullage through ventilation. 

5.2.4 Comparison of Calculations.

Equilibrium THC values for each test were calculated with both models using the start of test 

thermal conditions as model inputs, as these were shown to be fairly stable conditions. This data 

is shown in figure 48 along with the measured THC at the start of each test. Also shown on the 

chart are the corresponding temperatures that were used for model inputs. All CWT calculations 

were performed using an assumed 3% fuel load, except test 4 which used a fuel load of 25 

percent, as per the test plan. 

The FAR calculator, as expected, overpredicts the THC in the CWT, due to its assumption of 

isothermal fuel tank conditions with a given fuel temperature. Thus, this model provides 

consistently conservative estimates of the FAR in the CWT, with noticeably increased 

predictions as the fuel temperature is increased. Using the FAR calculator for the wing tank 

estimates however, at times, underpredicts the THC in the tank. This is due to the lower fuel 

temperature that is observed in the wing tank. Assuming an isothermal condition, using the fuel 

temperature, provides a low estimate because the increased skin and ullage temperatures provide 

for reduced condensation effects and increased vapor generation in the tank. 

Because the vapor generation model is currently configured to work solely with CWTs, wing 

tank calculations are not shown. Looking at the CWT data, however, one can see that this model 

provides a more accurate prediction of the THC than the FAR model. The ability of the user to 

input separate fuel and tank wall temperatures allow more accurate modeling of the 

thermodynamic processes in the tank. The highest variation from the measured values observed 

with the vapor generation model was just 0.38% propane. 
0.00%

0.50%

1.00%

1.50%

2.00%

2.50%

3.00%

3.50%

Test 0

Test 1

Test 2

Test 3

Test 4

Test 5Test 6

THC (% Propane)

CWT Measured

CWT-FAR Calculator

CWT-Vapor Generation Model

WT Measured

WT FAR Calculator

FIGURE 48. MEASURED AND COMPUTED EQUILIBRIUM THC VALUES 

Because the vapor generation model is currently configured to work solely with CWTs, wing 

tank calculations are not shown. Looking at the CWT data, however, one can see that this model 

provides a more accurate prediction of the THC than the FAR model. The ability of the user to 

input separate fuel and tank wall temperatures allow more accurate modeling of the 

thermodynamic processes in the tank. The highest variation from the measured values observed 

with the vapor generation model was just 0.38% propane. 

Figure 49 shows the measured and predicted CWT THC results from test, using the vapor 

generation model, with all systems running on the ground. The model was run using two 

different flashpoint fuels: 130° and 135°F. The actual flashpoint of the test fuel, as given in 

appendix I, was 131°F. As expected, the two different computed results closely bracket the 

measured data, thus providing confidence in the models ability to accurately track the THC 

evolution in the CWT. 

Figures 50 through 52 show the computed and measured results of several of the flight tests 

through the entire flight profiles. Each of the calculated results were computed using fuel 

compositions according to the measured flashpoints (appendix I), matched with Woodrow’s

compositions at the same flashpoint [11]. The pressure data from each flight test were modified 

to remove any fluctuations that were occurring in flight, because these were found to have 

adverse effects on the results. The overall trend of the computed data matches closely with the 

measured results in all three cases, with the exception of the descent portion of figure 51. The

magnitude of the THC at the cruise altitude, however, is overestimated in each case by the

model. It is believed that this is attributable either to differences in the ambient pressure

readings and the tank pressure due to vent flow or a reduction in the accuracy of the

mathematical relationships used at reduced pressures. This issue is currently being evaluated,

49
and changes to the model will be made as needed. The trends of figure 51, as mentioned, match 

quite closely through the entire flight, with the exception of descent. At this point, the measured 

data rises, while the calculated results decrease. In this case, the measured data appears to be

incorrect for an undetermined reason, as this trend was not observed in any of the other flights

and is not what would be expected. 

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Time (s)

THC (% Propane)

Measured

Computed - 130 FP

Computed - 135 FP

FIGURE 49. CENTER WING TANK THC MEASURED AND COMPUTED RESULTS FOR 

A GROUND TEST 

0

0.5

1

1.5

2

2.5

3

0 2000 4000 6000 8000 10000 12000

Time (s)

THC (% Propane)

0

20000

40000

60000

80000

100000

120000

Pressure (Pa)

Measured

Calculated

Pressure

FIGURE 50. MEASURED AND CALCULATED THC MEASUREMENTS FOR 

FLIGHT TEST 3 

50
0

0.5

1

1.5

2

2.5

3

0 2000 4000 6000 8000 10000 12000 14000 16000 18000

Time (s)

THC (

% Propane)

0

20000

40000

60000

80000

100000

120000

Pressure (Pa)

Measured

Calculated

Pressure

FIGURE 51. MEASURED AND CALCULATED THC MEASUREMENTS FOR 

FLIGHT TEST 4 

0.0000

0.5000

1.0000

1.5000

2.0000

2.5000

3.0000

0 5000 10000 15000 20000 25000

Time (s)

THC (% Propane)

0

20000

40000

60000

80000

100000

120000

Pressure (Pa)

Measured

Calculated

Pressure

FIGURE 52. MEASURED AND CALCULATED THC MEASUREMENTS FOR 

FLIGHT TEST 5 

51
52

6. SUMMARY OF FINDINGS. 

The results of the tests indicated that the inerting system operated as expected, and the dual-flow 

methodology allowed the system to maintain an ullage oxygen concentration below 12% for the 

flight tests performed. Any deviations in system performance from test to test could be 

explained by the difference in system warmup times on the day of the test. Using the variableflow methodology allowed a greater amount of nitrogen-enriched air to be generated on descent 

at a higher oxygen concentration with flow being as much as 50% greater for short periods of 

time, but it had no measurable effect on the resulting average ullage oxygen concentration after 

each test when compared to the dual-flow methodology. The higher flow rates did seem to help 

inert gas distribution by decreasing the worst bay oxygen concentration. The highest average 

ullage oxygen concentration observed on any flight test correlates directly with the worst bay 

oxygen concentration, illustrating the importance of maintaining a low average ullage oxygen 

concentration for good inert gas distribution. Oxygen diffusion between the bays of the tank was 

relatively rapid and showed that diffusion would reduce the oxygen concentration in bay 1 (worst 

bay) from 13 to 12 percent in approximately 1 hour with the air cycle machines running. 

Overnight dispersion of the ullage oxygen concentration was measured to be very small when 

considering other effects that could affect an increase in oxygen concentration overnight such as 

fuel transfer and vertical gradient diffusion. 

Flammability measurements from both the center wing tank and number 2 wing tank showed 

trends very similar to what was expected, based upon both experimental and computer model 

data. The equilibrium data agreed favorably with the data from both the Fuel Air Ratio 

Calculator and Condensation Model, while transient data also matched closely with that of the 

Condensation Model. The measurements generated in these flight tests are being used to 

enhance the capability of these existing flammability models and will be used in the future to 

further improve upon predictive flammability calculations. 

7. REFERENCES. 

1. “Fuel Tank Harmonization Working Group Final Report,” Aviation Rulemaking 

Advisory Committee, July 1998. 

2. Burns, Michael and Cavage, William M., “Inerting of a Vented Aircraft Fuel Tank Test 

Article With Nitrogen-Enriched Air,” FAA report DOT/FAA/AR-01/6, April 2001. 

3. Clodfelter, R.G., Anderson, C.L., and Vannice, W.L., “OBIGGS for Fighter Aircraft,” 

SAE Technical Paper 871903, Aerospace Technology Conference and Exposition, Long 

Beach, California, October 5-8, 1987. 

4. Burns, Michael and Cavage, William M., “Ground and Flight Testing of a Boeing 737 

Center Wing Fuel Tank Inerted With Nitrogen-Enriched Air,” FAA report 

DOT/FAA/AR-01/63, August 2001. 

5. Burns, Michael et. al., “Flight Testing of the FAA Onboard Inert Gas Generation System 

on an Airbus A320,” FAA report DOT/FAA/AR-03/58, June 2004. 
53/54

6. Summer, S., “Limiting Oxygen Concentration Required to Inert Jet Fuel Vapors Existing 

at Reduced Fuel Tank Pressures—Final Phase,” FAA report DOT/FAA/AR-04/8, August 

2004. 

7. Burns, Michael and Cavage, William M., “A Description and Analysis of the FAA 

Onboard Oxygen Analysis System,” FAA report DOT/FAA/AR-TN03/52, June 2003. 

8. Ochs, Robert and Polymeropoulos, C.E., “Vaporization of Jet A in a Simulated Aircraft 

Fuel Tank Including Sub-Atmosphere Pressures and Low Temperatures,” Proceedings 

from the International Fire and Cabin Safety Conference, Lisbon, Portugal, November 

15-18, 2004. 

9. Cavage, William M. and Kils, Ole, “Inerting of a Boeing 747SP Center Wing Tank Scale 

Model With Nitrogen-Enriched Air,” FAA report DOT/FAA/AR-02/51, May 2002. 

10. Cavage, William M., “Ground-Based Inerting of Commercial Transport Aircraft Fuel 

Tanks,” RTO-MP-103 26, Proceedings from RTO AVT Specialists’ Meeting on Fire 

Safety and Survivability, Aalborg, Denmark, September 23-26, 2002. 

11. Woodrow, J.E., “Flash Point and Chemical Composition of Aviation Kerosene (Jet A),” 

NTSB Docket SA-516, Exhibit No. 20S, May 2000. 
A-1/A-2

APPENDIX A—CENTER WING TANK INSTRUMENTATION PANEL

INSTALLATION DRAWING

APPENDIX B—HEATED GAS SAMPLE FITTINGS INSTALLATION DRAWING

B-1/B-2 

C-1/C-2

APPENDIX C—INSTRUMENTATION DIAGRAM WITH CHANNEL LISTING 
To Left

Surge Ta

4

OEA [O2]

Power Cable

Empty Pack Bay

Purge Air Line

System Wiring

Sample Tubing

Instrumentation Wiring

1

2

5

4

7

8

3

6

#2 Wing

 Tank

Modifed Fastener 

WS400

FAS

Air

OBOAS

Existing

Power

Rack

Window Plug

Air

Air

Air

OBOAS

Compressed Air Bank

(Cargo Bay)

Penetration STA 1275

Modified Waste

Tube, STA 1235

Anal

FAS

yzer

Sample 

System

OBIGGS

DAS

Harness #7&8Harness #5&6

Harness #TBD

Harness #TBD

Cabin Temperature

Sys Pressures

Ch. 59-63, 

CWT Temps

Ch. 19-42

Sys O2

Ch. 70-71

Sys O2 Pres

Ch. 62-64

System/Pack 

Bay Temps

Ch. 66-73

Pressure

Transducers

CWT Oxygen 

and THC 

Analyzers

Ch. 1-8

Ch. 43

OBOAS/FAS 

Pressures

Ch. 9-18, 44-45

OBOAS/FAS 

Temperatures

Ch. 46-48

Ejector Supply Shut-off

Sys Flow

Ch. 75

Misc Temps

Ch. 76-91

Wing Tank 

Temps

Ch. 51-58

H/X Bypass Shut-Off

Aircraft Cabin
Listing of Specified Data Acquisition Channels

Item Description Range Output Designator

Fuel Tank / OBOAS Instrumentation

1 Oxygen Channel - Bay 1 Right 0-25% 0-10 vdc

2 Oxygen Channel - Bay 1 Left 0-25% 0-10 vdc

3 Oxygen Channel - Bay 2 Right 0-25% 0-10 vdc

4 Oxygen Channel - Bay 2 Left 0-25% 0-10 vdc

5 Oxygen Channel - Bay 3 0-25% 0-10 vdc

6 Oxygen Channel - Bay 4 0-25% 0-10 vdc

7 Oxygen Channel - Bay 5 0-25% 0-10 vdc

8 Oxygen Channel - Bay 6 0-25% 0-10 vdc

9 Bay 1-R O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

10 Bay 1-L O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

11 Bay 2-R O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

12 Bay 2-L O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

13 Bay 3 O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

14 Bay 4 O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

15 Bay 5 O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

16 Bay 6 O2 Anylz Inlet Pressure 0-500 Torr 0-1 vdc

17 OBOAS 1 Outlet Pressure 0-1000 Torr 0-1 vdc

18 OBOAS 2 Outlet Pressure 0-1000 Torr 0-1 vdc

19 Ullage Temp - Bay 1 Right 0-450 F T-Type

20 Ullage Temp - Bay 1 Left 0-450 F T-Type

21 Ullage Temp - Bay 2 Right 0-450 F T-Type

22 Ullage Temp - Bay 2 Left 0-450 F T-Type

23 Ullage Temp - Bay 3 0-450 F T-Type

24 Ullage Temp - Bay 4 0-450 F T-Type

25 Ullage Temp - Bay 5 0-450 F T-Type

26 Ullage Temp - Bay 6 0-450 F T-Type

27 Fuel Temp - Bay 1 Sump Area 0-450 F T-Type

28 Fuel Temp - Bay 2 Middle 0-450 F T-Type

29 Floor Temp - Bay 1 Right 0-450 F T-Type

30 Floor Temp - Bay 1 Left 0-450 F T-Type

31 Floor Temp - Bay 2 Right 0-450 F T-Type

32 Floor Temp - Bay 2 Left 0-450 F T-Type

33 Floor Temp - Bay 3 0-450 F T-Type

34 Floor Temp - Bay 4 0-450 F T-Type

35 Floor Temp - Bay 5 0-450 F T-Type

36 Floor Temp - Bay 6 0-450 F T-Type

37 CWT Wall Temp - Forward Tank Wall 0-450 F T-Type

38 CWT Wall Temp - Left Tank Wall 0-450 F T-Type

39 CWT Wall Temp - Right Tank Wall 0-450 F T-Type

40 CWT Wall Temp - Left Tank Wall 0-450 F T-Type

41 CWT Ceiling Temp - Forward 0-450 F T-Type

42 CWT Ceiling Temp - Aft 0-450 F T-Type

43 THC - CWT (FAS1) 0-50K PPM 0-10 vdc

44 THC - Wing Tank (FAS2) 0-50K PPM 0-10 vdc

45 FAS 1 Inlet Pressure 0-500 Torr 0-1 vdc

46 FAS 2 Inlet Pressure 0-500 Torr 0-1 vdc
Item Description Range Output Designator

Fuel Tank / OBOAS Instrumentation (cont'd)

47 Spare Outlet Pressure 0-1000 Torr 0-1 vdc

48 OBOAS Temp 1 0-450 F T-Type

49 OBOAS Temp 2 0-450 F T-Type

50 FAS Temp 0-450 F T-Type

51 Wing Tank Forward Spar Surface Temp 0-450 F T-Type

52 Wing Tank Inboard Fuel Temp 0-450 F T-Type

53 Wing Tank Rear Spar Surface Temp 0-450 F T-Type

54 Wing Tank Ullage Temp 0-450 F T-Type

55 Wing Tank Mid Fuel/Ullage Temp 0-450 F T-Type

56 Wing Tank Outb’rd Wall Surface Temp 0-450 F T-Type

57 Wing Tank Bottom Surface Temp 0-450 F T-Type

58 Wing Tank Top Surface Temp 0-450 F T-Type

System Instrumentation

59 Pack Bay Static Pressure 0-15 psia 0-5 vdc

60 Bleed Pressure 0-100 psia 0-5 vdc

61 ASM Inlet Pressure 0-50 psia 0-5 vdc

62 ASM Exit Pressure 0-50 psia 0-5 vdc

63 Post Orifice Pressure 0-25 psia 0-5 vdc

64 Bleed Temperature 0-450 F T-Type

65 ASM Inlet Temperature 0-450 F T-Type

66 ASM Exit Temperature 0-450 F T-Type

67 Post Orifice Temperature 0-450 F T-Type

68 Heat Exchanger Inlet Temp 0-450 F T-Type

69 Heat Exchanger Outlet Temp 0-450 F T-Type

70 NEA Oxygen Concentration 0-25% 0-10 vdc

71 OEA Oxygen Concentration 0-100% 0-10 vdc

72 NEA O2 Analyzer Inlet Pressure 0-500 Torr 0-1 vdc

73 OEA O2 Analyzer Inlet Pressure 0-500 Torr 0-1 vdc

74 NEA/OEA Analyzer Outlet Pressure 0-1000 Torr 0-1 vdc

75 NEA Flow Meter .8-80 SCFM .5-5 vdc

Additional

76 Pack Bay 1 Temp 0-450 F T-Type

77 Pack Bay 2 Temp 0-450 F T-Type

78 Pack Bay 3 Temp 0-450 F T-Type

79 OBIGGs Bay Temp 0-450 F T-Type

80 Bleed Duct Area Temp 1 0-450 F T-Type

81 Bleed Duct Area Temp 2 0-450 F T-Type

82 Bleed Duct Area Temp 3 0-450 F T-Type

83 Heated Line Temp 1 0-450 F T-Type

84 Heated Line Temp 2 0-450 F T-Type

85 Heated Line Temp 3 0-450 F T-Type

86 Heated Line Temp 4 0-450 F T-Type

87 Heated Line Temp 5 0-450 F T-Type

88 Heated Line Temp 6 0-450 F T-Type

89 Heated Line Temp 7 0-450 F T-Type

90 Heated Line Temp 8 0-450 F T-Type

91 Cabin Temp 0-450 F T-Type
D-1/D-2

APPENDIX D—SURFACE-MOUNTED THERMOCOUPLE DRAWING

E-1/E-2

APPENDIX E—ONBOARD INERT GAS GENERATION SYSTEM MECHANICAL

DRAWING WITH INTERFACE ILLUSTRATION
Interface 2:

Heat Exchanger 

Cooling Air Inlet

Interface 1:

Heat Exchanger 

Cooling Air 

Exhaust

Interface 3:

System Bleed 

Air Connection

Interface 4:

System NEA 

Output

Interface Description:

Interface 1: 4” Dia. Silicone Bellows with hose clamp

Interface 2: 4” Dia. Silicone Bellows with hose clamp

Interface 3: 2” Dia SS Pipe with captive tubing mate and Perry Seal

Interface 4: 1” Dia Al tube with Hydraflow 14J02 Sleve, Coupling, Seal, and Clamp Assy

Interface 5: 2” Dia. Silicone Bellows with hose clamp

Static Pressure 

Temperature 

Spare [O2] 

Static Pressure 

Temperature 

NEA [O2] 

Temperatu

Notes:

All Instrumentation Ports 1⁄4-Inch SwageLok with 1/8

th

-Inch Tubing A

All Thermocouples T-type 1/8

th

-Inch Probe through SwageLok Fittin

Static Pressure 

Temperature 

Attachment 1: FAA OBIGGS Engineering Drawing

Interface 5:

System OEA 

Output

Temperature

OEA [O2] 
APPENDIX F—ONBOARD INERT GAS GENERATION SYSTEM INERT/DIVERT

TUBING ASSEMBLY TOP DRAWING

F-1/F-2 
Install OBIGGS Deposit System with 

Isolation/Diversion Valve Combinatio
G-1/G-2

APPENDIX G—NITROGEN-ENRICHED AIR DEPOSIT NOZZLE INSTALLATION

DRAWING 

APPENDIX H—ONBOARD INERT GAS GENERATION SYSTEM WIRING DIAGRAM

H-1/H-2

APPENDIX I—TEST FUEL FLASHPOINT AND DISTILLATION DATA

I-1
747 Flight Test Flashpoint Data 

Fuel From

Flight Test 

No. 

Flashpoint 

Sample 1 

(°C) 

Flashpoint 

Sample 1 

(°F) 

Flashpoint 

Sample 2 

(°C) 

Flashpoint 

Sample 2 

(°F) 

1 55.0 131.0 52.8 127.0

2, 4, 5 59.5 139.1 57.3 135.1 

3 57.5 135.5 55.8 132.44

6 55.0 131.0 53.3 127.9

ASTM D 2887 Simulated Distillation for Fuel From Flight Test 1 

% Off BPoF BPoC % Off BPoF BPoC % Off BPoF BPoC 

IBP 235.4 113.0 34 380.8 193.8 68 431.1 221.7

1 262.0 127.8 35 382.0 194.5 69 433.0 222.8

2 274.7 134.8 36 382.9 194.9 70 435.3 224.1

3 284.0 140.0 37 383.7 195.4 71 437.1 225.1

4 295.6 146.5 38 384.8 196.0 72 438.7 225.9

5 301.9 149.9 39 386.4 196.9 73 440.4 226.9

6 310.9 155.0 40 388.2 197.9 74 442.2 227.9

7 315.2 157.3 41 389.8 198.8 75 443.8 228.8

8 318.5 159.2 42 391.2 199.6 76 446.0 230.0

9 323.2 161.8 43 392.4 200.2 77 448.5 231.4

10 326.1 163.4 44 393.6 200.9 78 451.1 232.8

11 329.0 165.0 45 394.9 201.6 79 452.6 233.7

12 331.0 166.1 46 396.3 202.4 80 453.7 234.3

13 334.5 168.1 47 397.8 203.2 81 455.1 235.1

14 339.6 170.9 48 399.5 204.2 82 457.7 236.5

15 341.5 171.9 49 401.5 205.3 83 460.4 238.0

16 342.8 172.6 50 403.5 206.4 84 464.0 240.0

17 344.8 173.8 51 404.9 207.1 85 467.8 242.1

18 348.5 175.9 52 406.2 207.9 86 471.2 244.0

19 351.3 177.4 53 407.6 208.7 87 474.1 245.6

20 353.3 178.5 54 409.1 209.5 88 477.5 247.5

21 355.1 179.5 55 411.0 210.6 89 480.7 249.3

22 356.7 180.4 56 413.0 211.6 90 484.8 251.6

23 358.2 181.2 57 414.8 212.7 91 486.8 252.7

24 360.6 182.5 58 417.1 214.0 92 490.2 254.6

25 363.7 184.3 59 418.5 214.7 93 496.3 257.9

26 365.7 185.4 60 419.6 215.3 94 502.4 261.3

27 367.0 186.1 61 420.3 215.7 95 507.6 264.2

28 368.4 186.9 62 421.4 216.3 96 515.8 268.8

29 370.2 187.9 63 423.3 217.4 97 521.7 272.0

30 372.3 189.0 64 424.9 218.3 98 533.8 278.8

31 374.6 190.3 65 426.3 219.1 99 551.5 288.6

32 376.7 191.5 66 428.1 220.1 FBP 573.8 301.0

33 379.0 192.8 67 429.8 221.0

I-2
ASTM D 2887 Simulated Distillation for Fuel From Flight Tests 2, 4, and 5 

% Off BPoF BPoC % Off BPoF BPoC % Off BPoF BPoC 

IBP 264.2 129.0 34 382.9 195.0 68 431.5 222.0

1 273.9 134.4 35 383.7 195.4 69 433.4 223.0

2 285.7 140.9 36 384.5 195.8 70 435.6 224.2

3 298.0 147.8 37 385.6 196.5 71 437.4 225.2

4 308.6 153.7 38 387.4 197.5 72 438.8 226.0

5 314.5 156.9 39 389.1 198.4 73 440.4 226.9

6 318.2 159.0 40 390.5 199.2 74 442.1 227.8

7 323.3 161.8 41 391.7 199.8 75 443.8 228.8

8 326.4 163.6 42 392.9 200.5 76 445.7 229.9

9 329.3 165.2 43 394.1 201.1 77 448.4 231.3

10 331.5 166.4 44 395.3 201.8 78 450.9 232.7

11 336.1 168.9 45 396.5 202.5 79 452.4 233.6

12 340.4 171.3 46 398.0 203.3 80 453.4 234.1

13 342.0 172.2 47 399.5 204.2 81 454.5 234.7

14 343.3 172.9 48 401.6 205.3 82 456.7 235.9

15 345.8 174.3 49 403.4 206.4 83 459.0 237.2

16 349.7 176.5 50 404.9 207.2 84 462.3 239.0

17 352.0 177.8 51 406.1 207.8 85 465.6 240.9

18 354.0 178.9 52 407.3 208.5 86 469.4 243.0

19 355.8 179.9 53 408.8 209.3 87 472.3 244.6

20 357.2 180.7 54 410.6 210.3 88 475.1 246.2

21 358.9 181.6 55 412.7 211.5 89 478.5 248.1

22 361.6 183.1 56 414.5 212.5 90 482.3 250.2

23 364.5 184.7 57 416.7 213.7 91 485.7 252.1

24 366.1 185.6 58 418.4 214.7 92 487.6 253.1

25 367.4 186.3 59 419.3 215.2 93 491.8 255.4

26 368.8 187.1 60 420.1 215.6 94 498.4 259.1

27 370.3 188.0 61 420.9 216.0 95 504.7 262.6

28 372.4 189.1 62 422.2 216.8 96 511.6 266.4

29 374.5 190.3 63 424.0 217.8 97 518.1 270.1

30 376.7 191.5 64 425.5 218.6 98 529.8 276.5

31 378.8 192.7 65 426.9 219.4 99 546.0 285.6

32 380.8 193.8 66 428.7 220.4 FBP 569.7 298.7

33 382.0 194.5 67 430.1 221.2 

I-3
ASTM D 2887 Simulated Distillation for Fuel From Flight Test 3 

% Off BPoF BPoC % Off BPoF BPoC % Off BPoF BPoC 

IBP 249.9 121.0 34 381.7 194.3 68 435.3 224.1

1 268.8 131.6 35 382.5 194.7 69 437.0 225.0

2 282.3 139.1 36 383.3 195.2 70 438.4 225.8

3 295.1 146.2 37 384.5 195.8 71 440.1 226.7

4 301.6 149.8 38 386.3 196.8 72 441.8 227.7

5 310.7 154.9 39 388.2 197.9 73 443.5 228.6

6 314.7 157.1 40 389.7 198.7 74 445.4 229.6

7 318.0 158.9 41 391.2 199.6 75 447.9 231.1

8 322.6 161.4 42 392.5 200.3 76 450.4 232.4

9 325.7 163.2 43 393.7 200.9 77 452.0 233.3

10 328.4 164.7 44 395.0 201.7 78 452.9 233.9

11 330.5 165.8 45 396.5 202.5 79 454.2 234.6

12 333.8 167.7 46 398.2 203.5 80 456.4 235.8

13 338.9 170.5 47 400.3 204.6 81 458.9 237.2

14 340.8 171.6 48 402.3 205.7 82 462.2 239.0

15 341.8 172.1 49 404.1 206.7 83 465.3 240.7

16 343.6 173.1 50 405.4 207.5 84 469.0 242.8

17 347.2 175.1 51 406.8 208.2 85 471.8 244.4

18 350.4 176.9 52 408.4 209.1 86 474.4 245.8

19 352.3 177.9 53 410.5 210.3 87 477.4 247.5

20 354.1 179.0 54 412.5 211.4 88 480.3 249.1

21 355.9 179.9 55 414.5 212.5 89 484.2 251.2

22 357.4 180.8 56 416.9 213.8 90 486.1 252.3

23 359.6 182.0 57 418.2 214.5 91 488.5 253.6

24 363.1 183.9 58 419.1 215.1 92 493.4 256.3

25 365.1 185.1 59 419.9 215.5 93 499.2 259.5

26 366.5 185.8 60 421.0 216.1 94 504.5 262.5

27 367.9 186.6 61 423.0 217.2 509.6 265.4

28 369.6 187.5 62 424.6 218.1 96 516.3 269.1

29 371.9 188.8 63 426.1 219.0 97 522.4 272.5

30 374.2 190.1 64 428.0 220.0 98 534.2 279.0

31 376.4 191.3 65 429.4 220.8 99 552.2 289.0

32 378.7 192.6 66 430.9 221.6 FBP 576.2 302.3

33 380.5 193.6 67 433.0 222.8 

95

I-4
ASTM D 2887 Simulated Distillation for Fuel From Flight Test 6 

% Off BPoF BPoC % Off BPoF BPoC % Off BPoF BPoC 

IBP 236.8 113.8 34 380.9 193.8 68 439.8 226.6

1 264.0 128.9 35 382.2 194.6 69 441.5 227.5

2 277.4 136.3 36 383.1 195.1 70 443.1 228.4

3 287.7 142.0 37 384.0 195.6 71 444.8 229.4

4 297.7 147.6 38 385.2 196.2 72 446.9 230.5

5 304.9 151.6 39 387.4 197.5 73 449.5 231.9

6 311.3 155.2 40 389.6 198.7 74 451.7 233.2

7 314.8 157.1 41 391.3 199.6 75 453.0 233.9

8 317.6 158.7 42 392.6 200.3 76 454.0 234.4

9 321.3 160.7 43 394.1 201.2 77 455.6 235.3

10 324.9 162.7 44 395.9 202.1 78 458.2 236.8

11 327.3 164.1 45 397.9 203.3 79 461.0 238.4

12 329.8 165.4 46 400.0 204.4 80 464.3 240.2

13 331.6 166.4 47 402.2 205.7 81 467.8 242.1

14 335.0 168.3 48 404.3 206.8 82 470.8 243.8

15 339.6 170.9 49 405.9 207.7 83 473.2 245.1

16 341.4 171.9 50 407.5 208.6 84 475.8 246.5

17 342.5 172.5 51 409.4 209.7 85 478.4 248.0

18 343.7 173.2 52 411.7 210.9 86 480.6 249.2

19 347.1 175.1 53 413.8 212.1 87 484.1 251.2

20 350.2 176.8 54 416.1 213.4 88 486.2 252.4

21 352.2 177.9 55 418.1 214.5 89 487.8 253.2

22 354.2 179.0 56 419.2 215.1 90 491.1 255.1

23 356.1 180.1 57 420.1 215.6 91 496.1 257.9

24 357.7 181.0 58 421.2 216.2 92 500.8 260.4

25 360.0 182.2 59 423.1 217.3 93 505.4 263.0

26 363.8 184.3 60 424.9 218.3 94 509.0 265.0

27 365.8 185.5 61 426.6 219.2 95 515.3 268.5

28 367.5 186.4 62 428.5 220.3 96 518.6 270.4

29 369.1 187.3 63 430.3 221.3 97 526.5 274.7

30 371.2 188.4 64 432.2 222.3 98 536.6 280.3

31 373.9 190.0 65 434.6 223.7 99 551.6 288.7

32 376.2 191.2 66 436.8 224.9 FBP 570.9 299.4

33 378.8 192.6 67 438.2 225.7 

I-5/I-6