ad a095 057 - dtic · this final report is submitted by the general electric company, aircraft...

AFWAL-TR-80-2092ESL-TR-80-46

AD A095 057EVALUATION OF FUEL CHARACTER EFFECTSON J79 SMOKELESS COMBUSTOR

GENERAL ELECTRIC COMPANYAIRCRAFT ENGINE 4 SIt4i SS GROUPTECHNOLOGY PROGRAMS ANDPERFORMANCE TECHNOLOGY DEPARTMENTCINCINNATI, OHIO 45215

November 1980

TECHNICAL REPORT AFWAL-TR-80-2092, ESL-TR-80-46 -

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Final Report for Period July 1979 - June 1980

Approved for public release; distribution unlimited.

S AERO PROPULSION LABORATORYAIR FORCE 1WRIGHT AERONAUTICAL LABORATORIESAIR FORCE SYSTEMS COMMANDWRIGHT-PATTERSON AIR FORCE BASE, OHIO 45433

81 2 17 041'A'u

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NOTICE

When Government drawings, specifications, or other data are used for any purposeother than in connection with a definitely related Government procurement operation,the United States Government thereby incurs no responsibility nor any obligationwhatsoever; and the fact that the government may have formulated, furnished, or inany way supplied the said drawings, specifications, or other data, is not to be re-garded by implication or otherwise as in any manner licensing the holder or anyother person or corporation, or conveying any rights or permission to manufactureuse, or sell any patented invention that may in any way be related thereto.

This report has been reviewed by the Office of Public Affairs (ASD/PA) and isreleasable to the National Technical Information Service (NTIS). At NTIS, it willbe available to the general public, including foreign nations.

This technical report has been reviewed and is approved for publication.

J REY"S STTUD ARTHUR V. CHURCHILLFuels Branch Chief, Fuels BranchFuels and Lubrication Division Fuels and Lubrication Division

FOR THE COMMANDER

ROBE RT D. SHE'RRILL, ChiefFuels and Lubrication DivisionAero Propulsion Laboratory

"If your address has changed, if you wish to be removed from our mailing list, orif the addressee is no longer employed by your organization please notify AFWAL/POSF,W-PAFB, OH 45433 to help us maintain a current mailing list".

Copies of this report should not be returned unless return is required by securityconsiderations, contractual obligations, or notice on a specific document.AIk FORCE/56780/11 February 1981 - 230

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Block 20 (Continued)

iAt high power operating conditions, fuel hydrogen content was found to be avery significant fuel property with respect to liner temperature, flameradiation, smoke, and NOx emission levels.

At idle and cruise operating conditions, CO and HC emission levels were foundto be dependent on both fuel hydrogen content and relative spray dropletsize.

At cold day ground start conditions lightoff correlated with the relativefuel droplet size.

Altitude relight limits at low flight Mach numbers were fuel dependent andalso correlated with the relative fuel droplet size.

Combustor liner life analyses, based on the test data, yielded relativelife predictions of 1.00, 0.93, 0.83, and 0.73 for fuel hydrogen contentsof 14.5, 14.0, 13.0, and 12.0 percent, respectively.

High temperature cyclic fuel nozzle fouling tests revealed significanteffects of fuel quality and operating temperature on nozzle life. Theresults correlated with laboratory thermal stability ratings of the fuels.

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SEUIYCASFCTO FTI AEm/ aaEt~d

PREFACE

This final report is submitted by the General Electric Company, AircraftEngine Group, Evendale, Ohio. The work was conducted under Contract No.F33615-79-C-2033. Dates of Research were November 26, 1979 through March 2,1980. Author submittal date is November 1980. Program sponsorship andguidance were provided by the Aero Propulsion Laboratory (AFWAL/POSF),Air Force Wright Aeronautical Laboratories, Air Force Systems Command, Wright-Patterson Air Force Base, Ohio under Project 3048, Task 05, and Work Unit 98.Jeffrey S. Stutrud was the government project engineer.

Supplemental funding and technical guidance were provided in the areaof gaseous emissions and smoke measurement and analysis by the EnvironmentalSciences Branch of the Environics Division in the Research and DevelopmentDirectorate of HQ Air Force Enginering and Services Center, located atTyndall Air Force Base, Florida. Ibis organization has been formerly re-ferred to as CEEDO or the Civil Engineering Center.

Test fuel analysis was provided by APL, the Monsanto Research Labo-ratory (under contract to AFWAL), and the Air Force Logistics CommandAerospace Fuels Laboratory (SFQLA). The cooperation of these organizationsis appreciated.

Key General Electric contributors to this program were: K.L. Murrell,Program Manager; D.W. Bahr, Technical Program Manager; C.C. Gleason, Princi-pal Investigator; M.W. Shayeson, Fuels; T.L. Oller, High Pressure CombustorTests; W.F. Menkhaus, Low Pressure Combustor and Fuel Nozzle Tests; and M.J.Kenworthy/H.L. Foltz, Combustor System Life Analyses.

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TABLE OF CONTENTS

Section Page

I. SUMMARY 1

Ii. INTRODUCTION 3

III. TEST FUEL DESCRIPTION 5

A. General Description 5B. Physical and Chemical Properties 5

IV. J79 ENGINE COMBUSTION SYSTEM DESCRIPTION 7

A. Overall Engine Description 7B. Combustion System Description 7C. Combustor Operating Conditions 13

V. APPARATUS AND PROCEDURES 18

A. Performance/Emission/Durability Tests 18

i. High Pressure Test Rig Description 212. High Pressure Test Rig Instrumentation 213, High Pressure Test Procedures, 39

B. Cold-Day Ground Start/Altitude Relight Tests 39

i. Low Pressure Test Rig Description 392. Cold-Day Ground Start Procedure 423. Altitude Relight Test Procedure 42

C. Fuel Nozzle Fouling Tests 44D. Test Fuel Handling Procedures 48E. Data Analysis ProcedurcA 49

I. Fuel Property Correlation Procedures 502. Combustor Life Prediction Procedures 503. Turbine Life Prediction Procedures 52

VI. RESULTS AND DISCUSSION 55

A. Experimental Test Results 55

I. CO and HC Emissions 552. NOx Emissions 593. Smoke Emissions 654. Carbon Deposition and Emission 715. Liner Temperature and Flame Radiation 776. Combustor Exit Profile and Pattern Factor 877. Cold-Day Ground Starting and Idle Stability 958. Altitude Relight 989. Fuel Nozzle Fouling 105

V

TABLE CF CONTENTS (Concluded)

Section Page

B. Engine Syotem Life Predictions 111

1. Combustion Systew Life Predictions 1112. Turbine Life Predictions 113

C. Comparison of Results 113

VII. CONCLUSIONS AND RECOMMENDATIONS 121

A. Conclusions 121B. Recommendations 121

APPENDIX A - Fuel Property Data 123

APPENDIX B - High Pressure Test Data 132

APPENDIX C - Carbon Depositicn/Evission Data 153

APPENDIX D - Low Pressure Test Data 160

APPENDIX E - Fuel Nozzle Fouling Test Data 172

APPENDIX F - Smoke Data Calculation 175

REFERENCES 177

vi

LIST OF ILLUSTRATIONS

Figure

1. General Electric J79 Turbojet Engine. 8

2. J79 Engine Combustion System, Exploded Pictorial View. 9

3. J79-17C Low Smoke, Long Life Combustor Assembly. 10

4. J79-17C Combustor Flowpath. 11

5. J79-17C Combustor Inner Liner Details. 12

6. J79-17C Fuel Nozzle Tip Details. 14

7. J79-17C Fuel Nozzle Flow Characteristics. is

8. J79 Engine Altitude Windmilling/Relight Requirement Map. 17

9. Side View of High Pressure J79 Combustor Test Rig. 22

10. Combustor InsLallation in High Pressure J79 Test Rig. 23

11. Combustor Exit Instrumentation Rake. 24

12. High Pressure Test Rig Exit Instrumentation. 26

13. General Electric Smoke Measurement Console. 28

14. General Electric Emissions Measurement Console (CAROL IV). 29

15. Particle Fractionating Sample Details. 30

16. Combustor Metal Temperature Measurement' Locations andTypical Measured Temperature Levels. 32

17. Inner Liner Instrumentation of 0" CWALF. 33

18. Inner Liner Temperature Instrumentation at 180* CWALF. 34

19. Dome Temperature Instrumentation. 35

20. Rear Liner Temperature Instrumentation. 36

21. Fuel Nozzle Temperature Instrumentation. 37

22. Optical Pyrometer Setup in High Pressure Test Rig. 38

vii

LIST OF ILLUSTRATIONS (Continued)

Figure Page

23. Low Pressure J79 Combustor Test Rig. 41

24. Fuel Nozzle Fouling Test Rig - Overall Installation. 45

25. Fuel Nozzle Fouling Test Rig - Details. 46

26. Fuel Nozzle Fouling Test Cycle. 47

27. Node Model for J79 Comb-istor Showing Heat TransferQuantities. 51

28. Fatigue Diagram for Hastelloy-X Sheet Stock. 53

29. Turbine Nozzle Diaphragm and Instrumented Location ViewFactors. 54

30. Effect of Operating Conditions on CO Emission Levels. 56

31. Effect of Selected Fuel Properties on CO Emission Levelsat Idle and Cruise Operating Conditions. 58

32. Effect of Fuel Hydrogen Co-tent and Spray Droplet Size

on CO Emissions Levels. 60

33. Variation of HC Emission Levels with CO Emission Levels,Idle and Cruise Operating Conditions. 61

34. Effect of Fuel Hydrogen Content and Spray Droplet Size onIdle HC Emission Levels. 63

35. Effect oi Operating Conditions on NOx Emission Levels. 64

36. Effect of Fuel Hydrogen Content on NOx Emission Levels. 67

37. Effect of Flame Temperature on NOx Emission Levels. 68

38. Effect of Operating Conditions on Smoke Emission Levels. 69

39. Effect of Fuel Hydrogen Content on Smoke Emission Levels 72

40. Effect of Fuel Naphthalene Content on Smoke Number Normal-ized for Hydrogen Content. 73

41. Pattern Factor Variation During Carbon Endurance Test. 75

vili

LIST OF ILLUSTRATIONS (Continued)

Figure Pare

42. Effect of Fuel Hydrogen Content on Carbon Emission Levelsat Cruise (Test Point 4). 78

43. Carbon Particle Size Distribution at Cruise Measured byCascade Impactor. 79

44. Typical Effect of Operating Conditions on Liner Tempera-ture Rise, Fuel 1A. 81

45. Effect of Installation and Film Flow Mode on Liner

Temperature Response. 83

46. Rear Liner Hot Spot Caused by Thermocouple Leads. 84

47. Spatial Variation in Rate of Change of Liner Temperaturewith Fuel Hydrogen Content. 85

48. Effect of Operating Conditions on Rate of Change of LinerTemperature with Fuel Hydrogen Content. 86

49. Effect of Fuel Naphthalenes Content on Liner TemperatureRise Normalized for Hydrogen Content. 88

50. Effect of Combustor Operating Conditions on flameRadiation. 89

51. Effect of Fuel Hydrogen Content on Flame Radiation. 90

52. Effect of Fuel Naphthalenes Content on Radiant Heat FluxNormalized for Hydrogen Content. 91

53. Effect of Operating Conditions on Pattern Factor. 92

54. Variation in Pattern Factor with Test Sequence. 93

55. Comparison of Combustor Exit Temperature Profiles. 94

56. Typical Ground Start Characteristics. 97

57. Effect of Fuel Atomization on Gro,'nd Start. 100

58. Effect of Fuel Atomization on Altitude Relight Limits. 104

59. Posttest Appearance of Fuel Nozzle Primary Slot Pieces,Fuel Nozzle Fouling Tests. 106

ix

1 m m''4a

LIST OF ILLUSTRATIONS (Concluded)

Figure Page

60. Typical Effect of Time on Performance of J79-17C FuelNozzle with Heated Fuel. 108

61. Correlation of J79-17C Fuel Nozzle Fouling Test Results. 110

62. Typical Panel 0 Temperature Distribution. 112

A-1. JFTOT Results on Retained Sample of Fuel No. 2A1, byUSAF. 131

C-i. Posttest Photograph of Liner After 24-Hour Test ofFuel IA. 154

C-2. Posttest Photograph of Liner After 24-Hour Test ofFuel 13A1. 155

c-3. Posttest Photograph of Fuel Nozzle After 24-Hour Test ofFuel IA. 156

C-4. Posttest Photograph of Fuel Nozzle After 24-Hour Test ofFuel 13AI. 157

F-1. Experimental Relationship Between Smoke Number and ExhaustGas Carbon Concentration. 176

x

LIST OF TABLES

Table Page

I. Test Fuel Properties. 6

2. J79-17C Engine Combustor Operating Conditions. 16

3. J79-17C Engine Combustor Test Parts List. 19

4. Combustor Pretest Flow Calibrations. 19

5. Fuel Nozzle Pretest Flow Calibrations. 20

6 Summary of Measured and Calculated Combustor Parametersin High Pressure Combustor Tests. 25

7. Combustor Metal Temperature Instrumentation, High Pres-sure Rig Tests. 31

8. High Pressure Test Point Schedule. 40

9. Low Pressure Test Point Schedule. 43

10. Summary of CO Emission Test Results. 57

11. Summary of HC Emission Test Results. 62

12. Summary of NOx Emission Test Results. 66

13. Summary of Smoke Emission Test Results. 70

14. Summary of Carbon Deposition After 24 Hour Tests. 74

15. Summary of Total Carbon Emission Test Results. 76

16. Summary of Peak Liner Temperature Results. 80

17. Chronology of Hardware Changes During High Pressure Test

Series. 96

18. Summary of Ground Start Test Results. 99

19. Summary of Idle Stability Test Results. 101

20. Summary of Altitude Relight Limits. 102

21. Summary of Altitude Pressure Blowout Limits. 103

22. Summary of J79-17C Fuel Nozzle Fouling Test Results. 109

xi

LIST OF TABLES (Continued)

Table Page

23. Comparison of Engine Combustor Design Features. 114

24. Comparison of Operating Condition and Fuel PropertyEffects. 116

25. Comparison of Fuel Effects on Exhaust EmissionCharacteristics. 117

26. Comparison of Fuel Effects on Liner Durability

Characteristics. 11i

27. Comparison of Hydrogen Content Effects on Predicted

Combustor Liner Life. 120

A-I. Detailed Test Fuel Property Data. 124

A-2. Hydrocarbon Type Analysis Comparison. 125

A-3. Test Fuel Combustion Properties. 126

A-4. Thermal Stability Rating of Test Fuels. 130

B-1. Basic High Pressure Test Data. 133

B-2. Supplementary High Pressure Test Data. 135

B-3. CO Emission Test Data Correlation. 137

B-4. HC Emission Test Data Correlation. 138

B-5. NOx Emission Test Data Correlation. 139

B-6. Smoke Emission Test Data Correlation. 140

B-7. Detailed Inner Liner Temperature Data. 141

B-8. Detailed Rear Liner, Transition Duct and Fuel Nozzle StemTemperature Data. 144

B-9. Liner Temperature Data Correlation at Takeoff Point 7. 147

B-10. Liner Temperature Correlation with Hydrogen Content. 148

B-11. Flame Radiation Data Correlation. 149

B-12. Combustor Exit Temperature Profile Data. 150

B-13. Pattern Factor Test Data Correlation. 152

xii

LIST OF TABLES (Concluded)

Table

C-I. Pattern Factor During 24 Hour Tests. 158

C-2. Detailed Cascade Impactor Carbon Particle Size Data. 159

D-1. Altitude Relight Test Results, Fuel Number 1A. 161

D-2. Altitude Relight Test Results, Fuel Numbers 2AI and 3A. 162

D-3. Altitude Relight Test Results, Fuel Numbers 4A and 5A. 163



D-6. Altitude Relight Test Results, Fuel Numbers 10A and IIA. 166

D-7. Altitude Relight Test Results, Fuel Numbers 12A and 13AI. 167

D-8. Ground Start Test Results, Fuels IA Through 3A. 168

D-9. Ground Start Test Results, Fuels 4A Through 7A. 169

D-10. Ground Start Test Results, Fuels 8A Through IA. 170

D-11. Ground Start Test Results, Fuels 12A and 13A1. 171

E-1. Fuel Nozzle Fouling Test Results. 173

xiii

NOMENCLATURE

Symbol Units

A Area cm2 , ram2

CO Carbon monoxide

CO2 Carbon dioxide

CWALF Clockwise aft looking forward ---

E1 Pollutant emission index g pollutant/kg fuel

H Fuel hydrogen content (mass fraction) %

HC Hydrocarbon (calculated as CH4) --

N Fuel naphthalene content (volume fraction) %

JFTOT Jet Fuel Thermal Oxidation Tester

NOx Total oxides of nitrogen (=NO+NO2 , calculated as NO2)

Q Heat of combustion (net) MJ/kg

S Combustor operating severity parameter

SMD Sauter mean diameter --

SN Smoke number (by ARP #1256)

T Temperature K

V Velocity m/s

W Mass flow rate g/s, kg/s

X Exhaust gas pollutant concentration mg pollutant/kg air

b Curve fit equation intercept ---

f Fuel/air ratio g fuel/kg air

h Absolute humidity g H20/kg air

k Arbitrary constant

xv

h1SCEDING PAE SLAM(.NOT FILM~

NOMENCLATURE (Concluded)

Symbol Units

m Curve fit equation slope

n Hydrogen-to-carbon atom ratio

Heat flux kW/m2

r Curve fit correlation coefficient

x Independent variable

y Dependent variable ---

AP Pressure drop mPa

aT Temperature rise K

n Combustion efficiency %

v Kinematic viscosity M2/s

P Density kg/m 3

a Surface tension mN/m

*(f) Fuel/air ratio function (Figure 35)

Subscripts

3 Compressor exit station (Combustor inlet)4 Combustor exit station8 Engine exit stationc Combustore Effective

f Fuelm Me-ared

r Reference

st StoichiometricL Liner (metal)

gs Gas sampleTC Thermal (Thermocouple)

s Sampleavg Average

max MaximumOGV Compressor outlet guide vaneTND Turbine nozzle diaphragm

xvi

SECTION I

SUMMARY

The purpose of this program was to determine by combustor rig tests anddata analyses, the effects of fuel property variations on the performance, ex-

haust emission and durability characteristics of the General Electric Jq-17C(low smoke, long life) engine combustion system, and compare the results tothose previously obtained in similar tests of the J79-17A (in use, high smoke)and FI01 (advanced low smoke) combustion systems. Thirteen refined and

blended fuels, which incorporated systematic variations in hydrogen content(11.9 to 14.5 weight percent) aromatic type (monocyclic or dicyclic), initialboiling point (298 to 409 K by gas chromatograph), final boiling point (554 to

646 K also by gas chromatograph), kinematic viscosity (0.90 to 3.27 mm2 /s at

294.3 K) and thermal stability breakpoint (518 to 598 K by JFTOT), were eval-uated in: (a) 14 high pressure/temperature combustor performance/emissions/durability tests; (b) 14 low pressure/temperature combustor cold-day ground

start/altitude relight tests; and, (c) 7 high temperature cyclic fuel nozzle

fouling tests.

At high engine power operating conditions, (takeoff and supersonic dash),fuel hydrogen content was found to be a very significant fuel property with

respect to smoke, oxides of nitrogen (NOx), liner temperature, and flame radia-

tion levels. Each of these parameters increased with decreasing fuel hydrogencontent. Dicyclic aromatics tended to cause somewhat higher smoke levels, but

no other discernable effect of any other fuel property wa5 found. Carbonmonoxide (CO) and unburned hydrocarbon (HC) emission levels were so low atthese operating conditions that no trend with fuel properties could be de-tected.

At engine ground idle and subsonic cruise conditions, the same strongeffects of fuel hydrogen content on smoke levels was evident, but NOx levelsbecame virtually independent of any fuel property. Emission levels of COand HC were found to be jointly dependent on fuel hydrogen content and rela-

tive fuel spray droplet size calculated from fuel viscosity, density and

surface tension.

Combustor exit temperature profile and pattern factor (in high pressure

tests) were essentially independent of fuel type, but a sensitivity to liner

thermocouple installations was found.

In cold day ground start tests (to 239 K), lightoff was obtained withall fuels. At standard day conditions, lightoff fuel/air ratio was indepen-

dent of fuel type, but at lower temperatures, the lightoff fuel/air ratioincreased with the less volatile more viscous fuels and correlated with therelative spray droplet size.

Altitude relight test results were similarly dependent upon ambient

temperature and fuel properties. At low flight Mach numbers, where fuel and

L,1

air temperatures are low, relight altitude limits correlated with the relativespray droplet size. At high flight Mach numbers where fuel and air tempera-ture are elevated, relight altitude limits were nearly independent of fueltype.

High temperature cyclic fuel nozzle tests with JP-8 and No. 2 dieselfuels revealed significant effects of fuel type and operating conditions onnozzle fouling rates. Nozzle life was correlated with fuel temperature andfuel breakpoint (by JFTOT).

Combustor liner life analyses, based on the test data, were conducted.These analyses resulted in relative life predictions of 1.00, 0.93, 0.83, and0.74 for fuel hydrogen contents of 14.5, 14.0, 13.0 and 12.0, respectively.Turbine system life is not predicted to change for any fuels with propertieswithin the matrix tested.

The fuel property effects observed in these tests of the J79-17C combus-tion system are generally similar to those previously determined for theJ79-17A (in use, high smoke) and FlOI (advanced, low smoke) combustionsystems, except for relative combustor liner life predictions. Because ofthe high front end cooling effectiveness, the J79-17C combustor life is pre-dicted to be significantly less sensitive to fuel hydrogen content than arethe other two combustor designs.

2

SECTION II

INTRODUCTION

For more than 25 years, the primary fuel for USAF gas-turbine-poweredaircraft has been JP-4, a wide-cut distillate with excellent combustion char-acteristics and low-temperature capability. Typically, its heating value hasbeen over 43.5 MJ/kg (18,700 Btu/lb), its freezing point below 219 K (-650 F),and its aromatic content quite low, around Il percent by volume. A prime con-sideration in the definition of JP-4 was that during wartime a large percent-age of domestic crude oil could be converted into this product with minimumdelay and minimum impact on other major users of petroleum products.

Conversion from high volatility JP-4 to lower voilatility JP-8, which issimilar to commercial Jet A-1, as the primary USAF aircraft turbine fuel hasbeen under consideration since 1968. The strong motives for the change areNATO standardization and reduced combat vulnerability.

Domestic crude oil production peaked in 1971 and has been steadily de-clining since that time, while demand has continued to increase. Thus,particularly since 1973, the cost and availability of high-grade aircraftturbine fuels have drastically changed. These considerations have spurredefforts to determine the extent to which current USAF fuel specifications canbe broadened to increase the yield from available petroleum crudes and,ultimately, to permit production from other sources such as coal, oil shale,and tar sands.

As a result of the current and projected fuel situation, the USAF hasestablished an aviation turvine fuel technology program to identify JP-4 and/or JP-8 fuel specificatiors which:

I. Allow usage of key world-wide resources to assure availability

2. Minimize the total cost of aircraft system operation

3. Avoid sacrifices of engine performance, flight safety, or environ-mental impact.

Engine, airframe, logistic, and fuel processing data are being acquired toestablish these specifications. This report contributes to the needed database by describing the effects of fuel property variations on the GeneralElectric J79-17C engine main combustion system with respect to performance,exhaust emissions, and durability. Similar programs, based on the GeneralElectric J79-17A and FI01 engines and the Detroit Diesel Allison TF41 andhigh Mach engines have been previously conducted (References 1, 2, 3, and 4).Collectively, these programs provide representative data for the engineclasses that are expected to be in substantial use by the USAF in the 1980's.

3

This report summarizes the results of a 9-month, 3-task program which wasconducted to identify which fuel properties are important to J79-17C (low-smoke, long life) engine combustor operation and quantitatively relate fuelproperty variations to combustor performance, emission characteristics, anddurability characteristics. Also, wherever possible, these results have beencompared to those previously obtained for the JP79-17A (high smoke, in-use)and F1O (advanced, low smoke) engine combustion systems in order to illus-

r trate how fuel property sensitivity is affected by combustor design featuresand/or engine cycles.

Thirteen test fuels provided by the USAF were utilized. Descriptions andproperties of these fuels are presented in Section III. In Task I of the pro-gram, test planning and preparations were made, based on use of the J79 enginecombustion system components and operating characteristics described in SectionIV, and on the three test rigs and procedures described in Section V. In TaskIl of the program, 35 tests (14 high pressure/temperature combustor performance/emissions/durability tests, 14 low pressure/temperature combustor cold-dayground start/altitude relight tests, and 7 high temperature fuel nozzle foul-ing tests) were conducted. These are summarized in Section VI-A. In Task IIIof the program these test data were analyzed to establish the fuel propertycorrelations also presented in Section VI-A and to establish the engine systemlife predictions presented in Section VI-B. Finally, these results are com-pared in Section VI-C to those previously obtained, and conclusions and recom-mendations drawn from these tests and analyses are summarized in Section VII.

4

SECTION III

TEST FUEL DESCRIPTION

A. General Description

Thirteen test fuels were supplied by the USAF for combustion systemevaluation in this program. The fuels included a current JP-4, a current

JP-8, and a No. 2 diesel. The blends were made up by the USAF to achieve

three different levels of hydrogen content: 12, 13, and about 14 percentby weight. Two different types of aromatics were used to reduce the hydrogencontent of the base fuels: a monocyclic aromatic (xylene bottoms), and adicyclic aromatic described by the supplier as "2040 solvent" (a naphthaleneconcentrate). A third blend component, used to increase the final boilingpoint and the viscosity of two blends, is described as a Mineral Seal Oil,a predominantly (90 percent) paraffinic white oil.

The rationale for the selection of this test fuel matrix was to span sys-tematically the possible future variations in key properties that might be dic-tated by availability, cost, the change from JP-4 to JP-8 as the prime USAFaviation turbine fuel, and the use of nonpetroleuin sources for jet fuel pro-duction. The No. 2 diesel was selected to approximate the Experimental Referee

Broad Specification (ERBS) aviation turbine fuel that evolved in the NASA-Lewis

workshop on Jet Aircraft Hydrocarbon Fuel Technology (Reference 5).

B. Physical and Chemical Properties

As before, all fuel property data were provided to General Electric bythe USAF from either their own in-house or contracted (Monsanto Research Cor-poration) laboratory tests. Key fuel properties are summarized in Table I,

and additional detailed data are presented in Appendix A.

5

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SECTION IV

J79 ENGINE AND COMBUSTION SYSTEM DESCRIPTION

A. Overall Engine Description

The 379 engine is a lightweight, high-thrust, axial-flow turbojet enginewith variable afterburner thrust. This engine was originally qualified in1956, aad since that time various models with improved life and thrust havebeen developed. The model currently in use by the USAF, the J79-17A, was thereference engine for the previous fuel character effects program (Reference 1).

A later model, the J79-17C, which incorporates a low-smoke long-life combustionsystem, is the reference engine for this fuel character effects program.

An overall view of the J79 engine is presented in Figure 1. The J79 hasa 17-stage compressor in which the inlet guide vanes and the first six statorstages are variable. The compressor pressure ratio is approximately 13.4:1.The combustion system is cannular with ten individual combustion liner assem-blies. The turbine has an air-cooled first-stage stator, and a three-stageuncooled rotor that is coupled directly to the compressor. The engine rotoris supported by three main bearings. The afterburner is fully modulating witha three-ring "V" gutter flameholder. Afterburner thrust modulation is accom-plished by means of fuel flow scheduling and actuation of the variable area,converging-diverging type exhaust nozzle.

B. Combustion System Description

The J79 engine employs a cannular combustion system which consists often individual combustion liner assemblies located between inner and outercombustion casings forming an annular passage. An exploded view of the systemwith the various components, including the compressor rear frame and the tur-bine first stage nozzle is shown in Figure 2.

A pictorial view of a combustion liner assembly is shown in Figure 3, anda flowpath is shown in Figure 4. Each combustor (or "can") consists of threeparts which are riveted together to form an assembly. The key feature of theiow-smoke long-life J79-17C combustor is a completely redesigned, shortened,inner liner shown in Figure 5. This part is a machined ring shell which hascontinuous film cooling slots in the cylindrical portion of the liner, and animpingement cooling manifold coupled with cooling slots in the dome transitionsection. Combustion air is introduced through a swirl cup, a secondary co-,wirler and bellmouthed thimbles (4 in ignition cans, 6 in non-ignition cans)in the dome transition region. The rear liner is a sheet metal shell havingpunched cooling louvers and dilution thimble holes. The outer liner is an air-flow guide to assure proper flow distribution to the inner and aft liners. Inan engine assembly, two of the combustors are provided with spark ignitors forstarting. Adjacent combustors are joined near the forward ends by cross igni-tion tubes to allow propagation of the flame from the combustors with a sparkignitor to the other combustorR. The liners are each positioned and held in

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NOZZLE SUPPORT

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COMBUSTION LINER COASION

COMPRESSOR REAR FRAME J

II '~.FIRST STAGE TUABIN.- NOZZLE

I~z~.:::OUTFR COMBUSTION CASING

CROSS-IGNITIONTUBE AND CLAMP

COMBUSTON LINE

Figure 2. J79 Engine Com~bustion System, Exploded Pictorial View.

9

bo

0 -

E00

P4

'-44

100

42

044

140Z

Impingement Cooling Cneto oln oeManifold

Double Barrel

Panel IPanel 5

Figure 5. J79-17C Combustor Inner Liner Details.

12

place by mounting bolts at the forward ends. Axial stack-up and thermal growthare accomplished by a sliding seal between the combustors and the transitionduct. The major liner material is Hastelloy X.

The transition duct provides a ring of ten oval inlet ports to accept theten combustion liners, and the exit of the transition duct is annular to matchup with the turbine flowpath. The transition section exit area is approxi-

mately one-half of the total exit area of the ten combustors providing an ac-celerating flow stream into the turbine. The transition duct is mounted byfive radial support pins in the inner combustion casing. These pins have asliding fit with the transition duct, allowing for differential thermal growth,but they maintain the axial location of the duct. Sliding seals are providedwith both the combustors and turbine.

All of the earlier J79 models featured the use of dual-orifice, pressure-atomizing type fuel nozzles, with a single fuel inlet and a fuel-flow dividingvalve external to the mounting flange. For the J79-17C model, the fuel nozzleis modified by replacement of the tip with one which has features to provideair-blast atomization, and improved fuel/air mixing. It also is longer to matewith the shortened inner liner. The J79-17C tip details are illustrated inFigure 6. The new tip has a central primary pressure atomizing fuel orificeand eight radial, low-pressure drop secondary fuel disributors on the fuelnozzle outside diameter. The fuel flow schedule for a single fuel nozzle isshown in Figure 7. The tip of each fuel nozzle is provided with an air shroudwhich directs cooling-air across the face and around the fuel orifices toreduce the tendency for carbon formation.

C. Combustor Operating Conditions

The combustor must operate over a wide range of fuel flow, inlet airflow,temperature and pressure. Table 2 presents the combustor operating parametersat four important steady-state engine conditions which are typically encoun-tered. At each of these conditions, fuel effects on combustor performance,emissions and life are of particular interest.

In addition to steady-state operation, the combustion system must pro-vide fQr starting over a wide range of conditions, ranging from cold dayground start to relight at high altitude. Figure 8 presents the altitudewindmilling/relight conditions.

S13

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MIL-F-7024

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Rel ight

0 1 2 3 4

Fuel Nozzle Pressure Drop, MPa

Figure 7. J79-17C Fuel Nozzle Flow Characteristics.

15

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,I obtained, butI some misses

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Figure 8. J79-17C Engine Altitude Windmilling/Relight Map.

17

_ y =j, =- -- - o

SECTION V

APPARATUS AND PROCEDURES

All combustor and fuel nozzle testing in this program was conducted inspecialized facilities at General Electric's Evendale, Ohio plant, using ap-paratus and procedures which are described in the following sections. Combus-tor performance/emissions/durability tests were conducted in a high pressure/temperature single-can combustor rig which is described in Section V-A. Inparallel, combustor cold-day ground start/altitude relight tests were con-ducted in a low pressure/temperature single can rig which is described inSection V-B. Also in parallel, high temperature fuel nozzle fouling testswere conducted in a small specialized test rig described in Section V-C.Special fuel handling procedures used in all of these tests are described inSection V-D. Finally, procedures employed in analyses of these data are de-scribed in Section V-E.

New, engine-quality, current-model J79-17C engine combustion system com-ponents, listed in Table 3, were utilized in these tests. Ptetest flow cali-brations of the combustors and fuel nozzles are presented in Tables 4 and 5.

A. Performance/Emission/Durability Tests

High pressure/temperature single-can-combustor rig tests were conductedat simulated J79 engine idle, cruise, takeoff, and dash operating conditionswith each of the fuels to determine the following characteristics:

i) Gaseous emissions (CO, HC, and NOx).

2) Smoke emissions.

3) Carbon particle emissions.

4) Carbon deposition.

5) Liner temperatures.

6) Flame radiation.

7) Combustor exit temperature profile and pattern factor.

8) Idle stability (lean blowout and ignition) limits.

Thus, a large part of the total data was obtained in these tests using appa-ratus and procedures described in the following sections.

18

Table 3. J79-17C Engine Combustor Test Parts List.

Part Name GE Part Number National Stock Number

Ignition Combustor Liner 7012M89G05 2840-00-126-2856Assembly

Fuel Nozzle 3031M36PO2 2925-00-126-5730(Parker-Hannif in 1445-623386)

Main Ignition Unit 106C5281P3 2925-00-992-7904L(Bendix 10-358765-1,115 VAC, 400 cps)

Main Igniter Plug 696D256P06 2925-00-126-2879(Champion FHE 281-3

Main Ignition Lea 106C5282PI 2925-00-065-0801(81482-59386)

Table 4. Combustor Pretest Flow Calibrations.

Combustor Effective Airflow Area, cm2

Low Pressure Test High Pressure TestCombustor S/N 03623 S/N 03551

Front Liner Assembly 30.02 30.20

Aft Thimbles and Trim Holes 29.75 29.41

Aft Cooling Louvers and Seal 9.97 10.83

Total 69.75 70.45

19

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20

1. High Pressure Test Rig Description

These tests were conducted in the Small Combustor Test Facility, Cell A5,located in Building 304 of the Evendale Plant. This test cell is equippedwith all of the ducting, fuel and air supplies, controls, and instrumentationrequired for conducting small combustor high pressure/temperature tests. Highpressure air is obtained from a central air supply system, and a gas-firedindirect air heater is located adjacent to the test cell. For the single-can-combustor rig, J79 engine idle, cruise, and takeoff operating conditions canbe exactly duplicated. Dash operating conditions can be exactly duplicatedwith respect to temperature, velocity, and fuel/air ratio, but pressure andflow rates must be reduced about 25 percent in order to be within the facilityairflow capability.

The High Pressure Combustor Test Rig, shown in Figures 9 and 10, exactlyduplicates a one-tenth sector of the engine flowpath from the compressor out-let guide vane (OGV) to the first-stage turbine nozzle diaphragm (TND). Asshown in Figure 9, the test rig inlet flange which bolts up to a plenum cham-ber in the test cell, incorporates a bellmouth transition to the simulatedOGV plane. The combustor housing is a ribbed, thick-walled vessel which formsthe inner flowpath contour and side walls, covered by a thick lid that formsthe outer flowpath contour. Figure 10 shows a combustor installed in thepressure vessel with the lid removed. The combustor exit engages an actualsegment of an engine transition duct. Immediately downstream of the transi-tion duct is an annular sector section which contains an array of water-cooledinstrumentation rakes in the TND plane, indicated by an arrow in Figure 9.Additional details of the exit instrumentation rakes are shown in Figure 11.Downstream of the rakes the combustion gases are water-quenched and the sectorflowpath transitions to circular, which is bolted up to a remote-operated back-pressure valve in the test cell ducting. Two other features of the test rigcan be seen in Figure 9: a flame radiation pyrometer, mounted to view thecombustor primary zone through a crossfire duct window; and bleed air pipesto withdraw, collect, and meter simulated turbine cooling airflow.

2. High Pressure Test Rig Instrumentation

A summary of the important combustor operating, performance, and emissionparameters which were measured or calculated in these tests is shown in Table 6.Airflow rates were measured with standard ASME orifices. Fuel flow rates weremeasured with calibrated turbine flowmeters corrected for the density and vis-cosity of each test fuel at the measured supply temperature. Combustor inlettemperature and pressure were measured with plenum chamber probes.

Combustor outlet temperature, pressure, and gas samples were measuredwith a fixed array of seven water-cooled rakes, arranged and hooked up asshown in Figure 12. Each rake contained five capped chromel-alumel thermo-couple probes, located on radial centers of area, and four impact pressure/gassample probes, located midway between thermocouple elements. As shown inFigure 12, eight of the impact probe elements were hooked up for total-pres-sure measurement, and the other 20 elements were manifolded to four heatedgas sample transfer lines leading out of the test cell to the gas composition

21

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Smoke/Gas Sample, 6 Elements Manifold to Line II

SParticulate Sample, 4 Elements Manifold to Line IIIRake-Element c Particulate Sample, 4 Elements Manifold to Line IV

AlFlD2 Smoke

B3 Console

D3G4 Gas Sample Each at

GI Sample Points 2-9GI Console

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Bl Sample at SelectedE3 Test Points forE4 Mass Emission Rate

and Particle SizeAnalyses

MiliporeSampler

El0A2 - IV

G3D4

Figure 12. High Pressure Test Rig Exit Instrumentation.

26

measurement instruments. Transfer Lines I and ii were connected through se-lector valves to a smoke measurement console (Figure 13) and a gas analysisconsole (Figure 14). Transfer lines III and IV were connected to heated, high-pressure, soot/particulate samplers located near the test rig. The lines thencontinued on to flow control/metering apparatus in the control room area. LineIII was connected to the filter holder (Millipore, 47 mm diameter glass fiberfilters) for total mass emission rate determination, and Line IV was connectedto an 8-'stage particle fractionating sampler (Anderson Sampler Inc. Mark 1Iiwith glass fiber filters) for particle size determination. Details of thefractionating sampler are shown in Figure 15.

The General Electric smoke measurement console (Figure 13) contains stan-dard test equipment which fully conforms to SAE ARP 1179 (Reference 6). Smokespot samples are collected on standard filter paper at the prescribed gas flowrate and at four soiling rates spanning the specified quoted soiling rate.The spot samples are later delivered to the data processing area, where thereflectances are measured and the SAE Smoke Number is calculated.

The gaseous emission measurement console, shown in Figure 14, is one ofseveral that were assembled to General Electric specifications for CAROL sys-tems (Contaminants Analyzed and Recorded On-Line) and that conform, generally,to SAE ARP 1256 (Reference 7). This system consists of four basic instruments:a flame ionization detector (Beckman Model 402) to measure total hydrocarbon(HC) concentrations, two nondispersive infrared analyzers (Beckman Models 865and 864) for measuring carbon monoxide (CO) and carbon dioxide (C02) concen-trations, and a heated chemiluminescent analyzer (Beckman Model 951) for mea-suring oxides of nitrogen (NO or NOx

= NO + NO2) concentrations. Each ofthe instruments are fully calibrated with certified gases before and aftereach test run; and periodically during testing, zero and span checks are made.Readings from the instruments are continuously recorded on strip charts andhand-logged on test and calibration points for later calculation of concentra-tion, fuel/air ratio, and emission indices, using a computer program which in-corporates the equations contained in ARP 1256. Gas sample validity waschecked by comparison of sample to metered fuel/air ratios.

Carbon deposition tendencies with JP-4 and Diesel No. 2 fuels wereassessed by conducting 24 hour tests with each fuel, starting with a cleancombustor and fuel nozzle. At the completion of these tests, the combustorand fuel nozzle was removed, visually inspected, and ptotographed to documeotthe carbon deposition tendencies.

Combustor metal temperature was measured by an array of 35 thermocoupleslocated on the assembly as listed in Table 7 and shown in Figures 16 through21. The inner liner instrumentation pattern was selected to provide detaileddata in the vicinity of the known hottest regions of the liner.

Flame radiation in the primary burning zone of the combustor was measuredby a total-radiation pyrometer (Brown Radiamatic, Type R-12), which can be seenin Figure 9. A diagram of the optical view path is shown in Figure 22. Thepyrometer sensing element is a thermopile which provides a direct currentvoltage output. The flame radiation is focused on the thermopile with a

27

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calcium fluoride lens which is transparent to radiation of wavelengths lessthan three microns. The pressure seal at the test rig wall is formed by a

sapphire window which is transparent to radiation of even longer wavelengths.The sapphire window was swept clean by a small flow of filtered air. The

pyrometer was calibrated by viewing a resistance-heated-carbon backbody fur-

nace through the same optical system used in the combustor test. The furnace

temperature was measured with a disappearing-filament optical pyrometer.

3. High Pressure Test Procedures

A total of 14 high pressure rig tests were run: one for each test fuel

plus a repeat test with fuel No. 1 to establish test variability. The same

fuel nozzle and combustor were utilized in all tests.

Each test was conducted to the 9-point test schedule shown in Table 8.

On Point No. 1, minimum lightoff and lean blowout limits at idle inlet condi-

tions were determined. On Points No. 2 through 9, steady-state operating,

performance, and emissions measurements were obtained at simulated engine

idle, cruise, takeoff, and dash operating conditions. At each of these simu-

lated engine operating conditions, data were recorded at two nominal fuel/air

ratios: 80 and 100 percent of the engine cycle value corrected for the tebt

fuel heating value. However, if the data indicated that the 100 percent fuel/

air ratio point would locally exceed the gas temperature limits of the exit

thermocouple rakes, a lower fuel/air ratio point was substituted. This limit

was usually exceeded at simulated takeoff and dash conditions, so lower fuel/

air ratio points were substituted and a higher fuel/air ratio was run at

cruise conditions.

B. Cold-Day Ground Start/Altitude Relight Tests

Low-pressure/temperature single can combustor rig tests were conducted at

simulated J79 engine ground cranking and altitude windmilling operating condi-

tions to determine tle cold-day ground start and altitude relight character-

istics of each of th test fuels. The apparatus and procedures which were

utilized are described in the following sections.

1. Low Pressure Test Rig Description

These tests were conduccud in the Building 301 Combustion Laboratory at

the Evendale Plant. This facility has capabilities for testing small combus-

tor rigs over a wide range of simulated ground start and altitude relight con-ditions. Liquid nitrogen heat exchangers are used to obtain low fuel and air

temperatures, and steam ejectors in the exhaust ducting are used to obtain low

combustor inlet pressures.

The low-pressure, single-can J79 combustor rig used in these tests is

shown in Figure 23. The combustor housing is made from actual engine parts

and the rig exactly duplicates a one-tenth segment of the engine combustion

system flowpath. Combustor inlet temperatures and pressures are measured with

probes in the plenum chamber. The combustor assembly is installed from therear of the combustor housing which bolts up to a segment of an engine combus-

tor transition duct. An array of thermocouples is located in the transition

39

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duct to sense ignition and blowout. This rig has no provisions for turbinecooling-air extraction.

Air obtained from the central supply system was dried at the facility toa dew point of about 240 K and metered with a standard ASME orifice. Fuelflow rates were measured with calibrated turbine meters corrected for thedensity and viscosity of each test fuel at the measured supply temperature.All temperature, pressure, and flow data were read on direct indicating in-struments (manometers, potentiometers, etc.) and hand logged by the test oper-ator.

2. Cold-Day Ground Start Procedure

The first part of the test with each fuel was structured to evaluatecold-day ground starting characteristics. The test point schedule is shownin Table 9. The airflow rate (0.318 kg/s) and combustor inlet pressi.re(ambient) were set to simulate typical engine ground starting conditions(1000 rpm). Fuel and air temperature were lowered from ambient to 239 K min-imum (JP-8 freeze point) in steps to simulate progressively colder days. Ateach temperature step, minimum ignition and lean blowout fuel flow rates weredetermined. Maximum fuel flow rate was taken as 12.6 g/s/can, which is wellabove the current engine minimum fuel flow rate (6.49 g/s/can). The testsequence was as follows:

1) With inlet conditions set, energize the igniter and slowly open thefuel control valve until lightoff is obtained. Record lightoff fuelflow rate. Deenergize igniter.

2) Slowly decrease fuel flow rate to blowout. Record lean blowout fuelflow rate.

3) Decrease fuel and air inlet temperatures in 5 to 8 K increments andrepeat Steps I and 2.

When the minimum temperature limit was established, the altitude relightportion of the test was initiated.

3. Altitude Relight Test Procedure

The second portion of the test with each fuel was structured to evaluatealtitude relight and stability characteristics. The test schedule is alsoshown in Table 9. Investigations were carried out at four airflow rates (0.23,0.41, 0.50, and 0.91 kg/s) selected to span the J79 engine altitude relightrequirement map (Figure 8). Air temperature was selected from the windmillingdata and ranged from 244 K to ambient. Fuel temperature was matched to theair temperature. The test sequence was structured to determine:

1) The maximum relight and blowout pressure altitudes with currentengine minimum fuel flow rates (4.22 g/s/can).

2) The minimum relight and lean blowout fuel flow rates at the relightaltitudes determined in (i).

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The test sequence was as follows:

1) With 15.? km altitude conditions set, energize the igniter, set fuelflow rate at 6.49 g/s, then increase combustor inlet pressure (re-duce altitude and flight Mach number) until ignition occurs. De-energize the igniter and record maximum relight altitude conditions.

2) With fuel flow rate at 6.49 g/s, slowly reduce combustor inlet pres-

sure until blowout occurs. Record maximum pressure altitude blowoutconditions.

3) Energize igniter and increase fuel flow until lightoff. Deenergizeigniter and record minimum lightoff fuel flow rate at maximum re-light altitude.

4) Slowly reduce fuel fluw rate until blowout. Record lean blowoutfuel flow rate at maximum relight altitude conditions.

5) Repeat Steps I through 4 at each airflow setting.

C. Fuel Nozzle Fouling Tests

Tests with the low and high thermal stability fuels (No. 2 Diesel andJP-8 respectively) were conducted to determine the relative tendency to causefuel nozzle fouling, which might be in the form of valve sticking and/ororifice plugging. The J79 fuel nozzle is known to have a long troublefreeservice life and to be quite tolerant of fuel property variations. It wasanticipated, therefore, that the test conditions would need to be far moresevere than encountered in normal service to produce significant fouling ina reasonably short time.

The tests were conducted in the Building 304 1/2 Combustion Laboratoryusing a 7.62 by 12.7 cm flame tunnel setup shown in Figures 24 and 25. Inthis setup, hot fuel is pumped through the fuel nozzle which is immersed ina high velocity hot gas stream to simulate engine operations. Thermocoupleswelded to the upstream side of the nozzle stem (Figure 21) were used to moni-tor and control the nominal peak metal temperature at 672 K which usuallyrequired an inlet gas temperature of about 720 K. Two thermocouples immersedin the fuel line close to the nozzle were used to monitor and control fuelinlet temperatures. A cylindrical fitting was fabricated and attached tothe nozzle tip to conduct spent fuel to a 3.785 m3 fuel cart.

Fuel from the supply system was heated co approximately 422 K by highpressure steam, then heated to the desired temperature using Therminol 55 heattransfer fluid. The latter was supplied by a Chromalox electric fluid heattransfer system, Model PFOV-650-9. Heat transfer from the Therminol to thefuel as well as from the steam to the fuel, were accomplished by GrahamHeliflow heat exchangers.

In each test, one-hour simulated mission cycles were run, as indicatedin Figure 26. For fifty-seven minutes of each hour, tests were run at asimulated steady-state cruise condition, but to reduce fuel requirements to

44

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a more manageable volume, the flow rate maintained was reduced to 13.4 g/s,which is only 40% of actual engine cruise flow rate. This is still highenough to assure flow through the secondary system. To add realism to thetest, flow rates were increased and decreased (as shown in Figure 26) withoutchanging heat inputs, to obtain flow divider valve action, short temperatureexcursions, and flushing action between the simulated cruise periods.

In order to avoid excessive temperatures in the fuel nozzle at startup

and shutdown, the following procedures were followed:

For startup:

1) Establish desired fuel flow rate.

2) Set desired fuel temperature.

3) Establish airflow rate of 0.45 kg/s.

4) Start burner and set desired stem temperature

For shutdown:

1) Turn off preburner fuel supply.

2) Turn off steam supply and Chromalox power supply.

3) Open Therminol bypass around Graham heat exchanger.

4) Remove insulation from nozzle.

5) When nozzle temperatures drop to 394 K, shut off fueland air supplies, remove nozzle and recalibrate.

Normally, calibration of the nozzle was accomplished at the start andafter every six hours of operation on the test cycle. However, if it appearedthat the calibration was, or might be, changing rapidly, recalibration was per-formed as frequently as every three hours. Under mild conditions which causedlittle or no change in calibration, the test was continued up to 75 hours.Under severe conditions, the test was terminated at around 20 hours, which wasadequate to establish a trend.

D. Test Fuel Handling Procedures

The test fuels were delivered as needed by a USAF multicompartment tanktrailer, and stored in General Electric multicompartment tank trailers andunderground tanks, depending on the availability of storage and the volumesof fuel required. All of this tankage was used previously only in cleandistillate service. Prior to loading with a test fuel, each tank was drainedand visually examined to be sure it was empty. It was then rinsed with asmall volume of test fuel and drained again before loading. In the case ofof underground tanks, they could not be completely emptied by their installedpumps because their suction lines terminated several inches above the bottom.

48

In these cases, the manhole covers were removed and a portable pump was usedto empty the tank. It was then rinsed with clean aviation kerosene and re-emptied. This process was repeated as often as necessary until the rinsingsappeared as clean as the clean kerosene.

The fuels for the low-pressure tests were handled in drums filled fromthe storage tanks. Before loading, the drums were examined to be sure theywere empty and clean. They too, were rinsed with a small volume of test fuelbefore loading.

All the storage tanks and drums of test fuels were carefully marked withthe appropriate fuel identification.

Before taking data on any test point, the fuel system connecting the testfuel supply tank with the test vehicle was flushed with a quantity of testfuel equal to at least twice the volume of the connecting system.

Each test site had a provision for taking a representative fuel samplefrom the fuel supply line as close as practical to the test vehicle. Dupli-cate 0.95 dm3 samples of each test fuel were taken during each run for alltests except the fuel nozzle fouling tests, in which case 3.8 dm3 sampleswere taken in epoxy-lined containers. Before taking any sample, the samplingline was first flushed, then the sample container was rinsed twice with thetest fuel. All sample containers were labeled for complete identification,logged, and delivered to the USAF for analysis.

It should be noted that the 13 test fuels used in this program wereblended anew to duplicate the 13 fuels used in the two preceding programs(References 1 and 2). Since it is likely that fuels of the same number,blended on two different occasions, will differ slightly in composition andanalysis, those used in this program have been distinguished by adding thesuffix "A" to their numerical designations.

Fuel No. 13 was delivered on two consecutive days, for a total quantityof 15.14 m3. This material was co-mingled and used in fouling tests Nos. 1,2, and 3, and depleted after test No. 3. A third shipment of 7.57 m3 of

Fuel No. 13 was then received, and used in fouling tests Nos. 4 and 6.

Since gas chromotographic analyses performed by the Air Force on the co-mingled shipments and on the third shipment showed a slight difference incomposition, it was believed that the two batches might have different thermalstability breakpoints. Therefore, they were treated as different fuels andtested separately on the JFTOT to determine their breakpoints.

E. Data Analysis Procedures

Generally standard data reduction and presentation techniques were em-ployed. Key parameters and calculation procedures are indicated in Table 6and Appendix B. Some additional special procedures are described in thefollowing sections.

49

I. Fuel Property Correlation Procedures

Analyses of the experimental test results were conducted to: (1) cor-relate the performance and emission parameters with combustor operating con-ditions; (2) as appropriate, correct the measured rig data to true standardday engine operating conditions; and (3) correlate the corrected data withthe appropriate fuel properties from Section II. Generally, these procedureswere identical to those utilized in the previous J79 combustor evaluations andare described in Reference i. In the few cases where new procedures have beenutilized they are described in the Results Section, Section VI.

2. Combustor Life Prediction Procedures

The advanced low-smoke J79-17C combustor has a much improved life comparedwith the older design, J79-17A, which was the subject of Reference I. The im-proved life is currently being demonstrated in U.S. Navy J79-10 service (Refer-ence 8) and the time between replacements is being extended as exnerience withthis combustor increases. The failure mode of most significance to this studyis the development of low cycle fatigue cracks in the forward portion of thecombustor due to thermal gradients. The cracks form in the vicinity of thecooling slots in the conical dome region and gradually propagate upstream.The conical shape of this structure, however, permits limited cracks to existwhile still retaining the structural shape, thus resulting in a life signifi-cantly longer than the time to crack initiation.

A second life limitation feature occurs in the vicinity of crossfireelements. However, this limitation is thought to be vibration induced andnot affected by fuel type.

The remainder of the liner, in general, lasts satisfactorily until hot sec-tion overhaul. However, since the cooling on these liners was not grossly overdesigned, if fuel type greatly increased the metal *Pmveratures, new lifelimiting regions could become important.

In its initial design and throughout its .evelopmen., /9-17C linertemperatures, stresses, and life have been calculated by ceta- computeranalyses, with adjustments to the heat transf-.r inputs as taL ta identifieathe magnitude of specific contributors to the liner hea ig. tL-.trring toFigure 27, the combustor is heated by convection an; radaLion from the hotcombustion gases. These gases are hottest in ', ,, in ,,nd of the burnerand drop toward the exit temperature as the a- - through the dilutionholes mixes and cools the gas. The local gas velocities and temperaturesare calculated by computer program based on the air distribution, and thesevalues are then modified as indicated by subsequent combustor test data. Thecombustor liner is protected by the film air introduced through the film cool-ing slot. The rate at which the hot combustor gases mix through this protec-tive film has been established from laboratory test data and combustor experi-ence for the various specific film slots throughout the liner. Additionalinputs to the heat transfer calculation include the flame radiation heating,metal radiation cooling, the convective cooling rates on the cold side of the

50

0

00

I-r

ocn

.,4

0H

cdJf .(0

0 . I0•

U (U

~ 0 $4.

5-' (Uo

41-

4.'

$ U .

C: o0

0 $

0

U 0

0 co -4 0

p I Hi

1.i 00

C3a EU -4

44 0 t

4 1 H) a,4uH 0

H j 00I 440)U0

E-4 E-45

r= rX4

0w 0

51

liner, and the impingement cooling rate on the cooling slot overhang. The

flame radiation being the least well-defined from other data can be determinedby back calculation from measured metal temperatures. With the aid of a com-puter program to calculate the thermal conduction through the metal structure,the above heat transfer inputs or correlations are used to calculate the de-tailed temperature distribution within the combustor liner. These tempera-tures are then used as input to a stress analysis program together with inputsfor mechanical and aerodynamic loads, which then calculates the elastic stressesthroughout th. structure. These stresses are then used together with low cyclefatigue material properties (Figure 28) to predict life to first cracking, withan appropriate multiplier from experience to determine total life.

3. Turbine Life Prediction Procedures

If alternate fuels created substantial changes in temperature pattern fac-tor or temperature profile in the combustor exit gases, changes in turbine com-ponent life would be predicted. However, as discussed in Section VI-A-6, nochanges were found in these parameters.

The turbine vane heat load is made up of both convection and radiation andthe relative levels vary around the perimeter of the vane. The leading edge ofthe vane has a view of the dome region and thus changes in flame luminositydue to fuel type would result in changes in the radiation load and correspond-ing changes in the vane leading edge temperature. The relative radiation loaddepends also on the view factor of the dome region. The rear liner and tran-

sition piece have small view factors of the dome and changes in these metaltemperatures, due to changes in fuel type, provide an indication of the vanetemperature changes. Figure 29 illustrates the very small view factor whichthe stator leading edge has of the luminous flame region and it also showsthe similar view factor for the instrumented liner location. Temperatureswere measured in the program on both the inner and rear liners as well as onthe transition piece as discussed in Section V-A-2. As discussed in SectionVI-A-6, the rear liner temperatures, which are well away from the high flameluminosity, were less affected by fuel changes than were the inner liner tem-peratures, and the effect decreased at locations farther downstream. A thermo-couple was located on the transition piece such that the view factof of thedome region was very similar to that of the stator vanes. The temperatureat this location was essentially unaffected by changes in fuel hydrogen con-tent. Since the transition piece temperature was unaffected by fuel hydrogencontent, negligible effects are predicted for the stator vane.

52

500 Temperature, K

11:: ___8103

2125

0

503

0

cJ

CC.

0

_

ciis.Id

a)

ccc

0 >

A co

II+

4J

ci 0

00~

54

SECTION VI

RESULTS AND DISCUSSION

All planned test series (35 total) were completed and no major problemswere encountered. In general, results were well ordered and consistent withpublished data insofar as comparisons could be made. Detailed test results,which are listed in Appendices B through E, are summarized and discussed inthe following section. Engine system life prediction analyses based on theseresults are then presented in Section VI-B. Comparisons of these and pre-viously obtained data are presented in Section VI-C.

A. Experimental Test Results

Fourteen high-pressure rig tests were conducted to obtain the performance/emission/durability data. These data are listed in Appendices B and C andsumwarized in Sections VI-A-l through VI-A-7. Fourteen low pressure rig testswere conducted in parallel to obtain the ground start and altitude relightdata. These data are listed in Appendix D and summarized in Sections VI-A-7and 8. Also in parallel, seven fuel nozzle fouling tests were conducted toobtain the data listed in Appendix E and summarized in Section VI-A-9.

i. CO and HC Emissions

Carbon monoxide (CO) and unburned hydrocarbons (HC) are both productsof incomplete combustion and are, therefore, generally highest at low poweroperating conditions (idle and cruise). Figure 30 shows the strong effect ofcombustor operating conditions on the CO emission levels with JP-4 fuel. At4dle operating conditions, the CO emission index is about 66 g/kg which corres-ponds to a combustion inefficiency of about 1.5%, and is about the same levelas that produced by the standard combustor (Reference i).

At cruise, takeoff, and dash operating conditions, the CO emission levelsare approximately 15, 2, and 1%, respectively, of the idle CO emission level,which indicates the strong effect of combustor inlet temperature and pressureon combustion reaction rates, and hence, on combustion efficiency and CO emis-sion levels. In this correlation, the fuel/air ratio and temperature effectswere determained by multiple regression curve-fit techniques of these data,and all are somewhat stronger than were found for the standard combustor(Reference I).

Carbon monoxide emission trends like those shown in Figure 30 were ob-tained with each of the fuels, and a summary listing of the results is pre-sented in Table 10. At takeoff and dash operating conditions, CO emissionlevels are low with all of the fuels, so no fuel property effect is evident.However, as shown in Figure 31, at both cruise and idle operating conditions,CO emission levels are clearly lowest with the baseline JP-4 fuel and up to50% higher with those fuels having reduced hydrogen content, reduced vapori-zation properties (as indicated by both 10 and 90% recovery temperatures) andreduced atomization properties (as indicated by calculated relative spray drop-let size, SMD/SMDjp_4 , from Table A-3).

55

i2102_i

- JP-4 Fuel (Fuel No. 1A)

0 Baseline Test, Reading 9-16

C] Repeat Test, Reading 41-48

Idleat10%

i 011

-. -

0 Takeoff Cruise

*Y-4

100 Dash

i0-1 2111 1 , L I 1 1 1 1 1 1 1 1102 10-1 100

S o 42 PA 1.25 (94 eP~ 3)J T/s. MPa, g/kg, K

A Figure 30. Effect of Operating Conditions on CO Emission Levels.

56

Table 10. Summary of CO Emission TestResults.

Fuel CO Emission Index, g/kg

Number Idle Cruise Takeoff Dash

[ IA 63.3 8.5 1.1 0.6

1AR 71.0 7.8 1.0 0.5

2AI 85.2 9.0 l.I 0.6

3A 100.4 10.1 0.9 0.4

4A 110.5 10.9 1.3 0.6

5A 92.1 10.1 1.2 0.6

6A 90.2 9.7 1.3 0.6

7A 104.2 11.6 1.2 0.6

8A 80.4 11.7 1.5 0.6

9A 87.6 11.3 1.6 0.6

10A 96.1 9.2 1.3 0.6

11A 96.1 9.2 1.2 0.6

12A 94.3 9.9 1.2 0.6

13A1 100.9 12.9 1.5 0.7

5

co )

+00 C9 0

0- 0 t. r

OC3 0

*I-0 O U

N D

n cl Q)

0 00

L- 0, 0t'Cd 3 Qr.in OW

Ii - 1 U) tn

0' 0.U

r4o 0003 0 1I n o -

E-4

00

0 r 4

000

C1)4

00

0 In I n 00 0 0 0 0

RNj/i 'X~pUJ UOTSSTWi3 03

58

The atomization properties (viscosity, density, and surface tension) andvaporization properties (recovery temperatures) of these fuels turn out to behighly interrelated, so it is not possible to separate their effects on COemission levels in these tests. The data do correlate somewhat better withrelative droplet size than either of the recovery temperatures. The relativeeffects of atomization and hydrogen content were therefore, determined bymultiplc regression analyses which are illlustrated in Figure 32. At bothidle and cruise operating conditions, statistically significant correlationswere obtained showing CO emission indices to be approximately inversely pro-portional to hydrogen content and directly proportional to the square root ofdroplet size. A similar joint dependency on fuel chemical composition andphysical properties was shown in References 1 and 9, but not in Reference 2.It, therefore, appears that the relative importance of the fuel chemical andphysical properties are quite configuration dependent.

Hydrocarbon emission levels generally have been found to follow the sametrends as do CO emissions, but to be more sensitive to combustor operatingconditions and to exhibit more variability. Both of these trends were observedin the present tests and are illustrated in Figure 33, where HC emission levelsare plotted against CO emission levels for the idle and cruise test points andall fuels. At idle the HC index is about 27 g/kg (2.3% inefficiency). Atcruise the levels are about two orders of magnitude lower, and at takeoff anddash conditions the levels were very low. There is considerable scatter inthe cruise and idle data, but the regression curve fit exponent (2.14) is ingood agreement with past experience (about 1.5 to 3.0). Table 11 summarizesthe HC results for all fuels and operating conditions. At idle operatingconditions, the HC emission levels are clearly lowest with the baseline JP-4fuel and up to 100% higher with those fuels having reduced hydrogen and/orreduced atomization evaporation properties. As with the CO data, a multipleregression analysis was made (Figure 34) to show the relative effects of atomi-zation and hydrogen content, and the results are similar; the HC emissionindex is approximately inversely proportional to hydrogen content and directlyproportional to the square root of the droplet size.

2. NO, Emissions

Oxide3 of nitrogen (NO.) may form from oxidation of nitrogen whichoriginated either in the air or in the fuel. Current jet engine fuels and allof the fuels used in this program contain negligible amounts of bound nitrogen,so the following discussion is applicable only to the "thermal" NOx productioncharacteristics of current and advanced fuels. Fuels containing significantquantities of bound nitrogen have been investigated in other programs, andtypical results are contained in References 9, 10, and 11.

In contrast to CO and HC emission, which are products of incomplete com-bustion and are, therefore, generally significant only at low power conditions,"thermal" NOx is an equilibrium product of high temperature combustion and is,therefore, highest at high power operating conditions. Figure 35 shows thestrong effect of combustor operating conditions on NOx emission levels withJP-4 fuel. In this correlation, the pressure, temperature, velocity, humidityeffects were taken from previous studies (Reference 1) and the fuel/air ratio

59

0 JP-4 Based Fuels

JP-8 Based Fuels

A No. 2 Diesel Fuel

Tails: Monocyclic20 __ Aromat ics

Cruise

15

A

100 1

Z7-:y = 8.5x

0.8 1.0 1.2 1.4 1.6

-0 92 0.45SSMD , Fuel Parameter for Cruise CO

5 120

0 Idle

100

80080 .. .. 0 - 1. ..

60 y'74.8

40

0.8 1.0 1.2 1.4 1.6

S09 ( )D , Fuel Parameter for Idle CO

Figure 32. Effec:t of Fuel Hydrogen Content and Spray Droplet

Size on CO Emission Levels.

60

102

1 0 0 5 0 .1 / X 2 .1 4

Co Symbol Fuel No.0 IA

wx C [2A

10o~1 3A.74A

5A3 6A

Q 7AQ 8A

9A

11A0 12A

O 0 13A

Sio-2 101..10

10 0 10 1 10 2 103

CO Emission Index, g/kg

Figure 33. Variation of HC Emission Levels with CO Emission Levels,Idle and Cruise Operating Conditions.

61

Table 11. Summary of HC Emission TestResults.

Fuel HC Emission Index, g/kg


IA 26.6 0 0 0

1A 27.2 0 0 0(Repeat)

2A1 35.5 0.5 0.1 0

3A 48.7 0.4 -0.1 0.2

4A 50.7 0.7 0.2 0.1

5A 37.1 0.2 0 0

6A 34.5 0.6 0.2 0.1

7A 47.5 0.6 0.1 0

8A 35.4 0.4 0.1 0.1

9A 47.5 1.2 0.1 0.1

10A 50.2 1.2 0.4 0.2

11A 38.8 0.2 0.1 0

12A 46.9 0.5 0.2 0.1

13A1 44.9 1.0 0.2 0

62

60

0 JP-4 Based Fuels

W C JP-8 Based FuelsMtkoNo. 2 Diesel '

x Fuel

0 VTails: Monocyclic040Aromaticsb

20 ____

0.8 1.0 1.2 1.4 1.6

H )'.19 /SD \0.5214.5~ (SMDJP-. 4 ) Fuel Parameter for Idle HC

Figure 34. Effect of Fuel Hydrogen Content and Spray Droplet Sizeon Idle HC Emission Levels.

63

30

JP-4 Fuel

Tails: Repeat Test Dash

when f -. 11.525 (f) - 0.1243-0.233f

when f > 11.50(f) 1.461-0.0231f

20

a- Takeoff

Ve w.

S150 Cu

0 Cruise I

z 10

y 12.5x

Idle

0 0.5 1.0 1.5 2.0 2.5

S _(28.6) ( 1-p (664 . 6.29-h3x .359) 192 53.2

Figure 35. Effect of Operating Conditions on NOx Emission Levels.

64

i

effect was determined by curve fitting the new data. At takeoff operatingconditions, the NOx emission index is about 12 g/kg. At dash, cruise andidle operating conditions, the NOx levels are approximately 180, 50 and 20%,respectively of the takeoff levels. These NOx emission levels are about20% higher than those produced by the standard high smoke combustor (Refer-ence 1), but are also in good agreement with data from J79 engines equippedwith low smoke combustion systems (Reference 12).

NDx results similar to those shown in Figure 35 were obtained in eachof the tests. Emission indices for each fuel are listed in Table 12 and areplotted against fuel hydrogen content in Figure 36. At idle and cruise opera-ting conditions, virtually no effect of fuel properties is evident, but at thehigh power operating conditions (takeoff and dash), NOx levels decreasedwith fuel hydrogen content. This dependence on fuel hydrogen content is ex-pecLtd in diffusion flame processes such as these because of the stoichiometricflame temperature dependence on fuel hydrogen content and, in turn, the strongeffect of flame temperature on NOX formation rates. Figure 37 shows theeffect of stoichiometric flame temperature (from Table A-3) on the NO, emis-sion levels at takeoff.

3. Smoke Emissions

Smoke, like CO and HC, is a product of incomplete combustion, but com-bustors with virtually 100% combustion efficiency can produce highly visibleexhaust plumes, because the soot particle sizes are of the same order as thevisible light wavelengths. The J79-17C combustion system produces nearly in-visible exhaust plumes when fueled with current hydrogen content fuels(Reference 13).

The effect of combustor operating conditions on smoke levels with JP-4fuel is shown in Figure 38. No simple operating parameter could be derivedfrom the data, so smoke number is merely plotted against conbustor fuel/airratio and keyed as to inlet conditions. Within the test range, there isvirtually no fuel/air ratio effect, and the repeatability and agreement withprevious engine measurements is fair. At true idle, cruise and 75% densitydash operating conditions, the smoke levels are approximately 18, 19, and 38%,respectively, of the smoke level at true takeoff operating conditions. Atfull density dash operating conditions, smoke levels might be expected to besomewhat higher at the combustor exit plane and then be partially consumed inthe afterburning process. Because of the uncertainty of the extent of thesetwo opposing processes, no corrections were made. However, in Figure 38 andTable 13, all of the data have been corrected to engine outlet fuel/air ratioaccording to the procedure outlined in Appendix F to account for turbine cool-ing air dilution of the main combustor products and allow comparisons to enginemeasurements. Also shown in Table 13 are corresponsing smoke emission indices,calculated from the smoke number by the procedure described in Appendix F.

65

r

Table 12. Summary of NOx Emission TestResults.

Fuel NOx Emission Index, g/kg()


1A 1.8 5.8 11.6 23.3

1A 1.9 6.6 12.0 23.7

(Repeat)2AI 3.7 6.3 11.7 23.6

3A 1.9 6.2 11.9 23.1

4A 1.3 6.2 13.1 23.5

5A 2.3 6.4 12.3 23.3

6A 2.6 --- 11.6 21.9

7A 1.8 6.4 12.0 23.3

8A 2.2 6.2 12.3 25.3

9A 2.0 5.6 12.9 27.2

10A 1.8 6.2 13.0 26.1

11A 2.9 6.2 ......

12A 2.1 6.0 11.8 23.7

13AI 1.5 6.6 13.2 26.3

(1)Corrected to ambient humidity of 6.3

g/kg

66

32

0024 i

0 -0.20

- Dash, y = 23.6 40 JP-4 Based Fuels

M [ JP-8 Based Fuels

16 No. 2 Diesel Fuel0I r Takeoff, y 11.8

02 C

a0z! -0.03

0 Cruise, y 6.2N ~14.-5

0+0.67

Idle, y =2.2

0

11 12 13 14 15

Fuel Hydrogen Content, Weight, Percent

Figure 36. Effect of Fuel Hydrogen Content on NOx Emission Levels.

67

16

* Standard Day Takeoff JP-4 Based Fuels

Operating Conditions JP-8 Based Fuels

P3 = 1.359 MPa No. 2 Diesel Fuel

h 14 * T3 = 664 K0 H3 = 6.3 g/kg

i0 0 [3

12

\ y 11.8 + 0.0391 (x - 2494)

10

2480 2490 2500 2510 2520 2530

Tst, Stoichiometric Flame Temperature, K

Figure 37. Effect of Flame Temperature on NOx Emission Levels.

68

! 30* JP-4 Fuel* Open Symbols - Current Rig Tests* Tailed Symbols - Repeat Test* Solid Symbols - Previous Engine Tests

--- Average of Four Rig Data Points at Each.Operating Condition

Z 20

.4

Takeoff

00

MIdle

-- <______ _ [ Dash

.?0 0 Cruise

0

8 12 16 20 24

fs, Sample Fuel/Air Ratio (At Combustor Exit)

Figure 38. Effect of Operating Conditions on Smoke Emission Levels.

69

Table 13. Summary of Smoke Emission Test Results.

Fuel SN8, Smoke Number at Engine Exit Els, Smoke Emission Index, g/kg

Number Idle Cruise Takeoff Dash Idle Cruise Takeoff Dash

IA 1.9 2.3 9.7 3.2 0.015 0.011 0.045 0.017

IAR 2.9 2.8 17.1 7.0 0.026 0.017 0.083 0.035

2A1 3.1 8.6 13.2 3.7 0.027 0.055 0.063 0.019

3A 2.8 11.5 18.3 4.2 0.024 0.075 0.090 0.022

4A 12.2 25.9 23.7 14.0 0.118 0.213 0.130 0.080

5A 3.0 14.4 22.3 7.4 0.025 0.096 0.118 0.041

6A 10.7 12.7 13.7 5.7 0.124 0.086 0.063 0.031

7A 6.9 19.5 27.7 8.8 0.062 0.143 0.168 0.049

8A 6.0 19.9 40.0 8.2 0.054 0.148 0.288 0.045

9A 8.4 11.3 24.0 7.6 0.077 0.075 0.137 0.041

10A 14.4 15.9 31.0 7.9 0.141 0.111 0.208 0.042

11A 7.1 5.1 20.2 1018 0.067 0.031 0.102 0.060

12A 2.0 3.5 15.1 4.4 0.019 0.021 0.072 0.023

13AI 3.8 18.2 26.7 5.5 0.031 0.130 0.155 0.029

70

In Figure 39, the effect of fuel hydrogen content on engine smoke numberis shown. At each power level, smoke number increases exponentially withdecreasing hydrogen content, but there is considerable scatter in the data,particularly at takeoff operating conditions where the smoke levels arehighest. The scatter does appear to be associated with fuel volatility oratomization characteristics, but there is some trend with aromatic type, orfuel naphthalene content.

In order to quantify the effect of aromatic type on smoke levels, thesmoke numbers at each power level were normalized for hydrogen content, byuse of the regression equations shown in Figure 39, and plotted against fuelnaphthalene content as shown in Figure 40. Generally, the hydrogen normalizedsmoke numbers are seen to be within the range 0.5 to 1.5, but there is noclear trend with fuel naphthalene content. The line and equation shown inFigure 40 are the result of a second regression analysis of all of the nor-malized smoke numbers with fuel naphthalene content as the independent variable.The slope of the line (0.00711) is a measure of the average effect of fuelnaphthalene content on smoke number of constant hydrogen content. A 10% byvolume increase in fuel naphthalene content can be expected to produce a7.1% increase in smoke number. For comparison, a 1% by weight increase in fuelhydrogen content can be expected to produce an 89, 96, 31 and 36% increase insmoke number at idle, cruise, takeoff and dash operating conditions, respec-tively.

4. Carbon Deposition and Emission

As discussed in Section V-A-2, two 24-hour high pressure combustor testswere conducted with procedures established to provide information on therelative carbon deposition tendencies of Fuel 1A (Repeat JP-4) and Fuel 13AI(diesel). Both tests began with clean combustors and fuel nozzles. Inletconditions were varied over the idle, cruise, takeoff, and dash operating con-ditions on an approximately equal time basis. At the completion of each test,visual assessments of the relative ca.boning tendencies were made which aresummarized in Table 14. Photographs, which are included in Appendix C, werealso made to fully document the carbon deposition tendencies.

As shown in Table 14 and the posttest photographs, some caLbon depositionwas found in both tests, with the greater accumulation occurring using Fuel 13A(diesel). It should be noted that carbon deposition was not considered ex-cessive for either fuel and no adverse effects due to carboning are antici-pated using either fuel. For example, Figure 41 shows no long term changes inpattern factor using either fuel. These detailed pattern factor data arepresented in Table C-1 in the Appendices.

Also as discussed in Section V-A-3, carbon emission data were obtained onselected fuels using a millipore filter for total carbon collection, and acascade impactor for total carbon collection and particle size distribution.For both measurements, samples were withdrawn through the exit emission rakesat isokinetic conditions for 40 minutes. Table 15 sunmarizes the total carbonemissions measured by the millipore filler and cascade impactor.

71

/ No. 2 Diesel Fuel0 JP-4 Based Fuels

[ JP-8 Based FuelsTails: Monocyclic

Aromatics

50

Takeoff Cruise(Dry)

40 . . ..1 -- 3.6

-13.9( 1

4.5).5

020

z 3000

4-)4

Idle Dash

20 y 2.(11 0 i _ _

013 15 11 13 15

Fuel Hydrogen Content, Weight Percent

Figure 39. Effect of Fuel Hydrogen Content on Smoke Emission Levels.

~72

3,0Sb m

Idle 2.1 -8.48 - Cruise 3.9 -8.9

2.5 Takeoff 13.9 -3.6

0 Dash 4.3 -4.1

0 2.0

r=,0

Z 1.0)

-4vS 1.5

-1 9 0.00711x0.50

ZI

10 1.5 20 25 30

, x _Fuel Napthalene Content, Volume percent

Figure 40. Effect of Fuel Naphthalene Content on Smoke Number

Normalized for Hydrogen Content.

73

Table 14. Summary of Carbon Deposition After 24 Hour Tests.

Fuel 1A (Repeat) Fuel 13A

Fuel Nozzle Irregular deposits to 0.1 mm Hard black, shiny, uniformhigh on face. Black stain deposit on face to 0.8 mmaround secondary fuel ports thick and to 0.3 mm around

secondary ports

Venturi Soft, easily removed deposit Hard black deposits to b mm

to 1 mm thick thick

Inner Liner Very light discoloration Dull to shiny black deposits

easily wiped away up to 0.6 mm thick

Aft Liner Same as inner liner Light dull to shiny blackdeposit less than 0.1 mm

74

0.6 I

Performance w Carbon Endurance DataData (Table C-i)

0 0.4 - Doc C-C~r 0 0

__A<)

Test Point

9 0.2 0 2 6O 3 7<C> 4 Q 8

s Fuel 1A (Repeat) A 5 - 9

0I

0.6

9 Fuel 13A

Performance N Carbon Endurance Data

Data (Table C-i)0.4 (Table B-i) I

U0

0 DO0

44 0.2-_ _

0A

0 4 8 12 16 20 24

Elapsed Time, hr

Figure 41. Pattern Factor Variation During Carbon Endurance Tests.

75

Table 15. Summary of Total Carbon EmissionTest Results.

Carbon Particle

Emission, mg carbon/kg gas

Fuel Test Point Millipore CascadeNumber Number Filter Impactor

1A 4 11.3

7 9.7 ---

2A 4 6.4 ---

4A 4 --- 19.3

6A 4 10.5 ---

10A 4 17.1 12.0

13A 4 5.4 ---

7 29.6 ---

76

Figure 42 presents this data versus fuel hydrogen content for thecruise test point (Point 4). A moderate trend of increased carbon emissionwith decreased hydrogen content is noted. However, considerable data scatteris present.

Figure 43 presents the particle mass distributions for two fuels asmeasured by the cascade impactor. The detailed impactor data are presentedin Table C-2. Little difference in particle distribution is evident betweenthese two fuels. In general, about 50% of the carbon emissions by weight areless than I micron in diameter and about 30% are greater than 10 microns indiameter. This constitutes basically a two tailed distribution with eitherextremely small particles (less than I micron) or relatively large particles(greater than 10 micron) present.

5. Liner Temperature and Flame Radiation

Liner temperature measurements were obtained in the high pressure com-bustor tests at the locations described in Section V-A-2. Detailed data arelisted in Appendix B and peak data are summarized in Table 16. The data arepresented as metal temperature rise above the inlet air temperature (T2-T3).Typical levels and spatial variations at takeoff operating conditions areshown in Figure 16, and effects of variations in operating conditions on selec-ted typical thermocouples are shown in Figure 44.

Generally, as shown in Figure 16, the rear liner and transition metaltemperatures are quite moderate and uniform. The highest measured metal tem-peratures are in the forward or inner liner. The hottest thermocouple (No. 11)is an imbedded installation in the conical dome section first cooling slotoverhang. The internal swirling and burning flows emerge from the swirl cupas a separated flow, and create intense heat transfer to the conical domesurface, as they app.oach the wall and either reattach or create secoridaryflows. The film protection provided by the gill holes can be disrupted de-pending on the exact location of the reattachment process and result ineffects that extend further downstream. There are non-axisymmetric circum-ferential geometry features such as dilution holes, the igniter, and thecrossfire tube that result in circumferential variations in the flow reattach-ment process. This in turn results in extreme circumferential variations inmetal temperature as seen in Figure 16. To counteract this intense heating,the conical dome region is provided with intense cooling mechanisms includingimpingement on the rold side, internal convective cooling and conduction fromthe film ccuiing or gill holes, as well as the film protection. These intensemechanisms provide adequate cooling to treat the hottest regions but resultin much cooler metal temperature levels at locations that are not sufferingthe reattachment and film disruption.

As shown in Figure 44, inner liner temperature rises (both mid-panel sur-tace mounted and cooling ring overhang imbedded thermocouple installations)tended to be dependent upon operating pressure and temperature, but indepen-dent of operating fuel/air ratio. The absence of a fuel/air ratio effect inthe dome may be associated with near stoichiometric hot gas flows scrubbingthe conical dome walls. Variations in overall fuel/air ratio may be accompanied

77

40

Open Symbols - Using Cascade Impactor

Closed Symbols - Using MilliporeX Filter

S30

0

S20

r= 2- 5 )0

-3.

o r = -0.52

$4 10

0 A

0

11 12 13 14 15

Fuel Hydrogen Content, Weight percent

Figure 42. Effect of Fuel Hydrogen Content on CarbonEmission Levels at Cruise (Test Point 4).

78

100

Fuel 10AFuel 4A

40

20

1

4 10Ci,

2 2

0.4

10 40 80 95 99

Cumulative Weight Percent Less Than Stated Size

Figure 43. Carbon Particle Size Distribution atCruise Measured by Cascade Impactor.

79

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81.

by burning rate differences and features in the separated and reattachingflow regions along the conical dome that permit these near stoichiometricflame temperatures to exist over a significant range of overall fuel/airratios. This unusual flow phenomena is also thought to result in some random-ness in the data in this dome reattachment region.

As further shown in Figure 44, rear liner and transition metal tempera-ture rises, tended to be directly proportional to operating fuel/air ratio,

but independent of operating pressure on temperature. This behavior is thought

to indicate the absence of any significant radiant heat transfer to theseregions, due to the small flame view factor.

Detailed analyses of the liner temperature data to identify fuel propertyeffects was complicated by factors which are illustrated in Figures 45 and 46.Thermocouple attrition was such that, after reading number 56, the entire arraywas replaced for the final seven fuel test series (designated combustor instal-lation No. 3). In spite of extreme care to attach the new thermocouples in thesame spot, some thermocouples responded quite differently. As shown in Figure45, thermcouple No. 12 consistantly responded lower and thermocouple No. 13responded higher in the last installation. Other thermocouples, such as No. Ialso illustrated in Figure 45, exhibited two response levels associated withcooling film flow mode (attached or separated), but excellent reproducibilitybetween installations. Also, as shown in Figure 46, the liner thermocoupleleadout bundle created a blockage which disrupted the rear liner cooling flowthereby invalidating thermocouple No. 25 completely, and perhaps No. 26 and 27partially. Because of these installation and metastable flow phenomena, fueleffect analyses were mane for individual thermocouples, rather than for peakand average data as was done in References I and 2.

The effect of fuel hydrogen content on liner temperatur rise was de-termined for each thermocouple and installation/film flow mode by regressionanalyses, as illustrated in Figure 45, for the takeoff operating conditions.These results are tabulated in Appendix B, Table B-9, and summarized in Figure47, where the slope of the regression equation (Kelvin increase in linertemperature rise per percent decrease in fuel hydrogen content) is plot-ted against spatial location. Clearly, the forward panels of the inner linerare most sensitive (20 to 40 K/percent H) while the rear liner and transitionare virtually independent of fuel hydrogen content.

The effect of engine operating conditions on sensitivity of liner tem-perature rise to fuel hydrogen content was also determined by regression

analyses, for selected thermocouples. The trends are not easily generalized.One example is tabulated in Appendix B, Table B-10 and summarized in Figure 48.The thermocouple illustrated shows the largest sensitivity at cruise (35.2 K/percent H) and a relatively low sensitivity at takeoff (13.0 K/percent H),but other thermocouples respond differently. This Lhermocouple was selectedfor illustration to examine other fuel property effects, and make comparisonsto previous studies (in a following section).

82

0 JP-4 baseA Fuels

o JP-8 Based Fuels6 No. 2 Diesel Fuel

Tails: Nonocyclic Aromatics

Thermocouple No. 12 Takeolf Conditions

too (Panel I Midapen at 30')

(Point No. 7)

y - 359 + 12.35 (14.b - H)

400 Installation 1 & 2

200 j-00 Installation 3

I I I -

y 183 * 11.0 (14.5 H)

80

Thermocouple No. 13

(Panel 1 Overhang at 30*)

6 600

y 325 . 28.41 (14.5 - H)4o _ -I

V I I,.,.n.A 400

* i Installation 3

20 N) installation 1 &2

0 i I

800

Thermocouple No. 11

(Panel 0 Overhang at 30*)

600 I

Separated Film

400

0 Attached Film

200

(6 Fuels) I'I + 7 7.88 (14.5 H )

, , II I ,

11 12 13 14 15

Fuel Hydrogen Content, Weight percent

Figure 45. Effects of Installation and Film Flow

Mode on Liner Temperature Response.

83

VNV

~A

4=8,2

0

~Circumferent ial Location,40 Installation [] Degrees, CWALF&4 0 020

30 - 30 A Other

20 -0Solid: Imbedded on Cooling

lap Ring Overhang

S10

0

>-10

U

C:. 50Wr Installation 3.1 40

a0000

2 20 -

10 00w

E -10 E

Figure 47. Spatial Variation in Rate of Change of Liner Temperature with

Fuel Hydrogen Content.

85

o -Bae U

O JP-4 Based Fuels0 JP-8 Based Fuelsq

A No. 2 Diesel FuelTails: Monocyclic Aromatics

425

375

325

Dash(75 [P3) y 306.6 4 25.9 (14.5 -x)

275 I

375 I

I - y = 304.9 + 13.0o / (14.5- x)

Y;13.8 352-(4.5 - x

i Takeoff

$I 0 !l "e II /-y 147.8 + 35.2 (14. 5 -x)

!-:" 200

150

-ruise

100- -__________ _

150

I

y I 80.5 + 9.0 (14.5 -

-- --- - I' " -_i-__ > i - iZO - , t- "-- --..I A -.. ) _ -'

11 1213141

Fuel Hydrogen Cbnfent, Weight percent

Figure 48. Effect of Operating Conditions on Rate of Change of LinerITemperature with Fuel 4ydrogen Content.

86

As shown in both Figures 45 and 48, liner temperature rise does notappear to be dependent upon fuel atomization or vaporization properties,particularly at high temperature operating conditions. Also, the data donot appear to show sensitivity to aromatic structure, but this analysisis complicated because all of the 2040 solvent blends were tested in the thirdinstallation. However, to illustrate the effect of aromatic stuctures, thedata for thermocouple No. 7 (Figure 48) were noimalized for fuel hydrogencontent and plotted agoinst fuel naphthalene cortent in Figure 49 with theregression analysis line. As expeozed, this analysis shows a very low sen-sitivity of liner temperature rise to fuel naphthalene content at constantfuel hydrogen content.

Flame radiation measurements at the crossfire port plane were obtainedwith the apparatus and procedures described in Section VI-A-2; detailed re-sults are listed in Appdndix B. Due to an undetected noise source in thepyrometer signal, valid flame radiation measurements were not obtained forthe first six fuels tested. For the last seven fuels, Figure 50 shows theeffect of combustor operating conditions on flame radiation. The severityparameter shown was identified using regression analysis procedures. Theparameter is similar to the one derived for the J79-17A (R3ference 1) ex-cept no fuel/air ratio effect is evident in the current data. Figure 51,which summarizes the data of Table B-10, illustrates the effect of fu,,l hy-drogen content on flame radiation. Unlike the J79-17A high smoke com'.ustor,only a moderate fuel effect is observed. The absolute levels of flameradiation for the J79-17C combustor for all fuels are close to the level forJ79-17A combustor with JP-4 fuel, except at dash where the J79-17C valucsare about 50% of the J79-17A levels. These relatively low and fuel inselisi-tive levels may be due to the low smoke levels of the J79-17C combustorwhich would produce a nearly nonluminous flame for all fuels. In addition,the burning in the leaner dome J79--i7C combustor may occur largely upscreamof the crossfire port location and the pyrometer would, therefore, not viewpeak flame radiation but some lower radiation level from a cooler gas tem-perature. This latter possibility may also explain why some inner linertemperatures vary more with fuel hydrogen content than do the measured radi-ation.

Figure 52 presents radiant heat flux normalized for fuel hydrogen contentplotted against fuel naphthalene content to examine the effect of fuel aro-matic structure. The plot clearly shows that in this test series, dicyclicaromatics produce no greater ra-liant heat flux beyond that predicted by fuelhydrogen content.

6. Combustor Exit Profile and Pattern Factor

Combustor exit temperature distributions were measured in the high pres-sure tests using the fixed thermocouple rake artay described in Section V-A-2.Detailed datu are listed in Appendix B, and trends are illustrated in Fie,, -es53, 54, and 52

87

4j

eI I' I I* Thermocouple No. 7

(Representative Peak Surface Mounted Thermocoupi)*..~o-~b mta,, 2.0 _

2 . Idle 80.5 9.014 CS Cruise 147.8 35.244 $4 Takeoff 304.9 13.00 O

-4 Dash 306.6 25.95 ) 0.

0 1.0 IMP4 Z

V;- y =0.974 + 0.0043x

0 10 20 30

1 'I

Fuel Napthalenes Content, Volume Percent

Figure 49. Effect of Fuel Napthalenes Content on LinerTemperature Rise Normalized for Hydrogen Content.

88

C-4

00-4H

000

oo

'00

-1

I 0n

F4

Mt

-4 fn 0

o 4 uii4

012L

00

;:4 '0 W

cc Ln 0 ninC14~~ H x 1 -

a4:ss i :U n J I'l U T U4:

890

0 JP-4 Based Fuels

JP-8 Based Fuels

Tails: Monocyclic Aromatics

Idle Cruise

140 0

-y = 8 . 0.0 -Y = 094.8 + 1.32

(14.5 - H) 1 (14.5 - H) kj,

100 -

60I

1404

yi 120.9 +__ _3 11

60 lll 140 151135

Fuel Hydrogen C,-ntent, Weight percent

A

Figure 51. Effect of Fuel Hydrogen Content on Flame Radiation.

9/

100 1.y =.10.9....2.9

b m

Idle 83.2 0.60

Cruise 94.8 1.32Takeoff 106.9 2.05

Dash 120.9 2.93

N 2.5 T

1 .0

te y 1.015 -0.00137x

Vo 1 5

O 1.o ---_._

II" 0.5

0 10 20 30

Fuel Napthalenes Content, Volume Percent

Figure 52. Effect of Fuel Napthalenes Content on RadiantHeat Flux Normalized for Hydrogen Content.

91

0.6 r

<C. 1976 Test, JP-4 Fuel, Combustor S/N 405

l 1st Test This Series, JP-8 Fuel-- Combustor S/N 3551

0 2nd Test This Series, JP-4 Fuel

000.4

0

0,

Dash

Takeoff Cruise lule

0.9 1.0 1.1 1.2 1.3

-0.25 -0.50(AT ) We/T/P 3) Pattern Factor Operating Parameter

Figure 53. Effect of Operating Conditions on Pattern Factor. I

92

92i

Li

0 JP-4 Based Fuels

0 JP-8 Based Fuels

No. 2 Diesel FuelInstallation Tails: Monocyclic Aromatics

Number I la 2 3

0.40000 0

0.2 ,

Dash I0I

0.4

0 00

0.2 iTakeoff i

0 I I0

U(d 0I_ _ _ _ __

0.4

4

0.2 I

Cruise I0I

0.5

E0 dI 0 00 []0.3 I0 Z-- -- 0X

Idle

0.1 , I L.., L, , , ,

-4 -4 -4

Test Fuel Number (In Test Sequence)

Figure 54. Variation in Pattern Factor with Test Sequence.

93

I_

I

100

80 \ -__--___--

Takeoff Conditions

Open: Avg. Profiles0 Solid: Peak Profiles

060 -C 1976 Test, JP-4 Fuel - " I--4' Non-InstrumentedA Liner - S/N 405

00' This Study, Instrumented 4

Liner - S/N 3551

0 Rdg. 6, Fuel 2AZ ist Liner Installation

40'40) Rdg. 62, Fuel 10A

3rd Liner Installation

_ _ _ _ _ II/I I

20

/1 1

0.6 0.8 1.0 1.2 1.4

(T4 Local T3 )/(T4 Avg T3 ), Temperature Rise Ratio

Figure 55. Comparison of Combustor Exit Temperature Profiles.

94

It was anticipated from previous experience with this combustor modelthat exit temperature distributions would be relatively insensitive to com-bustor operating conditions and fuel type. As shown in Figure 53, both ofthese expected data trends were verified in the first two fuel evaluations,but the pattern factor levels were significantly higher than had been measuredin previous tests of the same combustor model, in the same test rig, as partof a vendor substantiation program. The differences in pattern factor levelswere too large to be attributed to either normal variations between combustorassemblies or test repeatabilities. But, at the time, no other obvious ex-planations could be found, so the test series were therefore continued asplanned.

In subsequent fuel evaluations, exit temperature distributions variedconsiderably, as shown in Figures 54, and 55, and a probable reason for thevariation was identified. In Figure 54, pattern factor (from Table B-12) foreach operating condition and fuel type is plotted in chronological test order,separated by "installation number" which is defined in Table 17. It is seenthat pattern factor levels were relatively constant within an installation,but varied significantly between installations. The pattern factor levels ininstallation No. 3 are in fair agreement with those expected from previcustests. As shown in Figure 55, detailed profiles also tend to display similarpeak levels, but noticeable difference in profile shape.

In Table 17, all hardware changes in the test seri.es which might haveaffected either combustor exit temperature distributions or temperaturemeasurements are listed. The major changes in pattern factor levels (Figure54) are associated with the two liner removals. Exit rake replacement andfuel nozzle seem to have had little effect on pattern factor levels.

A mechanism by which liner removal might affect pattern factor wasidentified in a posttest inspection of the outboard side of the rear liner.As shown in Figure 46, the liner thermocouple leadout blockage disruptedthe airflow to the rear liner in this region, causing a hot spot. Veryprobably the dilution hole airflows were also reduced, which could veryeasily have increased pattern factor. The effective blockage of the thermo-couple leadout wires probably varied between installations, thus introducingan extraneous variation in the test series. The most valid pattern factormeasurements, therefore, would be obtained with a noninstrumented liner, aswas used in the previous vendor substantiation program.

As noted above and shown in Figure 54, seven fuels were tested in in-stallation No.3, and pattern factor levels were relatively constant. JP-4and JP-8 based fuels with hydrogen contents of 12, 13, and 14 weight percentwere included in this installation series, so pattern factor must be vir-tually independent of fuel type.

7. Cold-Day Ground Starting and Idle Stability

Fourteen cold-day ground start tests were conducted in the low pressurerig using procedures described in Section V-B-2. Detailed test results arelisted in Appendix D and typical results are illustrated in Figure 56. In

95

AI

Table 17. Chronology of Hardware ChangesDuring Test Series.

Hardware Change Made After Run

FuelNozzle Liner

Installation Run Fuel Exit Removed for Removed Liner Re-No. No. No. Rakes Cleaning for Photos instrumented

1 I 2A,IA 2 rakes No No No

3 6A,5A,1IA No No No No

la 4 IAR No Yes Yes No

2 5 13A 4 rakes Yes Yes Yes

3 6 lOA,12A, No No No No9A,8A

7 4A,7A,3A End of This Test Series1 ... ..... i 9

93j

- - .~= - ----------. -- ~ -~-

lr - ~ -

C)

4.)

00

.1-4

00 u 00 r

C:4) 0 N - - C4 t

, 4 U)4

0 v

00

r-4 00C:10 C- 0

01 0

c00 C) ,-

PH 00 * __ _ 0 In

0

ICIA

0C 0

'H20 HT/T

'4.7

each test, 1000 rpm engine motoring conditions were simulated, and lean light-off and lean blowout limits were determined as a function of ambient (fueland air) temperature in steps from test cell ambient down to 239 K (-30 ° F).As shown in Figure 56, even with the most viscous fuel (No. 13AI, Diesel Fuel)lightoffs were obtaincd down to 239 K, but the fuel/air ratios required werefuel-type and temperature dependent. Lean blowout fuel/air ratios wereusually about half of those required for lightoff, with only minor fuel-type

and temperature dependence. Results for all fuels are summarized in Table 18.Effects of fuel properties on lean lightoff fuel/air ratio at both standardand cold day start conditions are illustrated in Figure 57. The standard daydata exhibit virtually no fuel dependence, but the cold day data are fuel de-pendent and correlate very well with the relative spray droplet size parameter(calculated from fuel viscosity, density and surface tension) which is derivedand listed in Appendix A.

Lean lightoff and blowout limits were also measured at idle operating con-ditions for all of the test fuels, as part of the high pressure test series.These data are summarized in Table 19. Lean blowout fuel/air ratios were lessthan 2.5 g/kg with all fuels. Lightoff fuel/air ratios ranged from 5.1 to 13.7g/kg. The variation is attributed to data scatter rather than to any fuelproperty effect. At these inlet conditions, both lightoff and blowout areprobably more dependent on fuel spray geometry and dome airflow than upon anyof the fuel properties.

8. Altitude Relight

Fourteen altitude relight tests were conducted in the low pressure rig

using procedures described in Section V-B-2. Detailed results are listed inAppendix D and summarized in Tables 20 and 21. In each test, altitude relight/blowout limits and lean relight/blowout limits were determined at four airflow/temperature levels selected to span the engine windmilling/air start map(Figure 8).

Altitude relight limits (Table 20) agreed well with previous rig testresults wherever comparison could be made, but tended to be pessimistic rela-tive to engine results (Figure 8) with respect to both altitude and flightMach number limits. Pressure blowout limits (Table 21) tended to be rela-tively close to relight limits in all cases.

Effects of fuel properties on altitude relight limits are illustrated inFigure 58. Like the ground start data, the altitude relight limits correlatewell with the relative spray droplet size parameter when the air and fueltemperatures are low (low Mach number flight conditions). At higher Machnumber conditions, where air and fuel temperatures are elevated, altituderelight limits become independent of fuel type. These fuel effect trendson both altitude relight and cold day ground start are very much like thosefound for the F101 combustor (Reference 2).

98

Table 18. Summary of Ground Start Test Results.(1 )

Fuel-Air Ratio, g/kg

Standard Day (288.2K) Cold Day (2 ) (239.0K)

Fuel Lean Lean Lean Lean

Number Lightoff Blowout Lightoff Blowout

1A 6.7 2.4 7.0 3.6

1A 6.5 2.5 8.0 3.7(Repeat)

2A 6.0 4.2 10.4 8.3

3A 6.8 4.5 23.9 10.3

4A 8.2 4.7 14.5 7.3

5A 6.2 4.3 9.3 7.3

6A 6.0 3.5 8.6 6.9

7A 7.5 5.0 11.9 8.8

8A 7.0 2.3 7.0 5.4

9A 6.5 2.0 9.2 5.4

10A 6.7 2.7 7.2 4.0

11A 6.2 2.5 6.5 4.1

12A 6.0 2.8 7.2 4.4

13A1 8.7 7.2 37.0 8.5

(1) Simulated 1000 rpm Cranking Conditions

P = 101 kPa

W - 3.18 kg/s per enginec

(2)All fuels light-off to 239K (at least)

99

A

7I

60

Standard Day (T3 Tf 288.2 K)

1000 rpm Cranking ConditionsP 101 kPa3

W f 3.18 kg/sec-Engine

40 ,.._

20

.1- 0 ________,,_5

S60 ,IL Cold Day (T Tf 239.0 K)

o 0 JP-4 Fuelsto,H C JP-8 Fuels/ DF-2 Fuel

w 40

20

0.8 1.0 1.2 1.4 1.6 1.8

SMD/SMDjp_4 , Relative Spray Droplet Size

Figure 57. Effect of Fuel Atomization on Ground Start.

100

Table 19. Summary of Idle Stability Test Results.

P 0.254 MPa

T -421K3

V - 24.2 m/s

Lightoff Lean Blowout

Wf, f WfV

Fuel Fuel Flow, Fuel-Air Fuel Flow, Fuel-AirNumber g/s/can Ratio, g/kg g/s/can Ratio, g/kg

1A 11.7 7.5 <2.3 <1.5

1A 11.6 7.4 <2.3 <1.5Repeat)

2A 12.1 8.0 <2.5 <1.7

3A 10.0 6.0 2.3 1.4

4A 18.3 11.8 3.8 2.5

5A 8.1 5.1 <2.5 <1.7

6A 10.8 7.1 <2.5 <1.7

7A 13.1 8.4 2.8 1.8

8A 21.8 13.? <1.3 <0.8

9A 16.8 10.2 11.6 7.7

10A 20.5 13.3 <1.3 <0.8

IIA 11.7 7.6 <2.5 <1.7

12A 11.7 7.6 <1.3 <0.8

13AI 10.0 6.7 <2.5 <1.7

101Z1

_ J=

Table 20. Summary of Altitude Relight Limits.

Wf 64.9 g/s

Wc 2.27 kg/b We 4.08 kg/s WC .99 kg/s W = 9.07 kg/sFuel Ait Alt Alt AltNumber km Mo km km Mo km Mo

IA 6.7 0.43 4.6 0.55 5.2 0.65 No Light

IA 7.3 0.46 6.1 0.62 6.1 0.69 No Light(Repeat)

2A 4.3 0.36 5.9 0.61 4.6 0.62 No Light

3A 5.2 0.39 5.8 0.61 6.3 0.70 3.1(1) 0.78(0)

4A 2.4 0.33 5.3 0.58 4.5 0.62 No ighc

5A 6.4 0.42 5.8 0.60 4.6 0.62 No Light

6A 6.0 0.41 5.6 0.60 4.0 0.60 No Light

7A 5.0 0.38 5.7 0.60 6.3 0.70 No Light

8A 6.6 U.43 7.1 0.37 5.7 0.67 No Light

I9A 6.6 0.43 7.9 0.70 6.3 0.70 No LightA

IOA 6.2 0.42 6.4 0.62 6.3 0.70 3.6(1) 0.76(1)

11A 6.7 0.44 6.9 0.66 6.8 1 0.73 4.9 0.84

12A 6.4 0.43 7.0 0.66 6.3 0.70 No Light

13AI 4.6 0.37 3.9 0.53 6.5 0.71 No Light

( 1)we 7.9 kg/s

102

Table 21. Summary of Altitude Pressure Blowout Limits.

Wf = 64.9 g/s

W =227 kg/s W =4.08 kg/s W 499 kg/s Wc =9.07 kg/sFuel Alt At Alt Alt

Number km M km Mo km Mo km M

1A 7.8 0.48 8.7 0.74 6.7 0.72 Not Determined

1A 8.4 0.51 8.4 0.73 8.0 0.78 Not Determined(Repeat)

2A 6.9 0.44 7.5 0.69 7.3 0.75 Not Determined

3A 7.7 0.48 7.5 0.68 7.8 0.77 9.3(1) 1.00(1)

4A 7.3 0.46 7.6 0.69 7.6 0.76 Not DeterminedI5A 6.8 0.44 7.9 0.70 8.0 0.78 Not Determined

6A 8.3 0.51 7.2 0.67 7.1 0.74 Not Determined

7A 7.0 0.45 7.9 0.70 6.5 0.78 Not Determined

8A 9.1 0.55 8.0 0.71 7.8 0.77 Not Determined

9A 8.3 0.50 8.9 0.75 7.5 0.75 Not Determined

I0A 8.2 0.50 8.1 0.71 8.6 0.81 8.8() 0.97(1)

11A 7.7 0.48 7.9 0.70 8.2 0.79 13.7 1.19

12A 7.8 0.48 8.4 0.73 7.5 0.76 Not Determined

13A1 5,,8 0.40 6.4 0.63 7.8 0.77 Not Determined

(1)W c 7.9 kg/s

I10

~103

-E--- - -.- -

JP-4 FuelsJP-8 Fuels Wf = 64.9 g/sec-EngineDF-2 Fuel

y 6.3-0.48x

50Z 0 _

M 0.70

Wc = 4.99 kg/sec

Tf =T3 - 289 K

100____ o__! 0

,0M = 0.60 1N

100

Wc = 4.08 kg/sec

Tf = T3 = 278 K

101

Iii

. ---y --10.6-3.84x

: -- Wc = 2.27 kg/sec" "

Tf = TI3 =244 K

0.8 1.0 1.2 1.4 1.6 1.8

SMD/SMDjp4 Relative Spray Droplet Size

Figure 58. Effect of Fuel Atomization on Altitude Relight Limits.

: 104

9. Fuel Nozzle Fouling

Seven fuel nozzle fouling tests were conducted, for a total of 346 hours.The first four and the sixth used Fuel No. 13, the low thermal stability No. 2diesel fuel. The fifth and seventh (Test No. 15) used Fuel No. 2, the highthermal stability JP-8.

The first test was run at 464 K, a temperature estimated to cause a moder-ate nozzle failure time, based on previous experience and knowledge of the fuelquality. However, the nozzle failed in only 14 hours and the test was termi-nated.

The second test was, therefore, run at a lower temperature of 450 K. Atthis condition, the nozzle showed only a small performance loss in 71 hours.

The third test was run at a higher temperature, 469 K, resulting in a mod-erate loss in performance in 54 hours.

The fourth test was run at a much higher temperature, 491 K, and a largeloss in performance was observed in 30 hours, though not as large as the lossin the first test at a much lower temperature.

A review of the first four tests at this time indicated that the firsttest was inconsistent with the other three. The tip of the first nozzle was,therefore, opened for inspection. The interior was found to contain heavy"carbon" deposits which would explain its performance, but which would not beexpected with the fuel temperature maintained. It was then observed that thenozzle stem was discolored, quite different from the other three, indicatingit had reached temperatures much higher than the desired 672 K gas streamtemperature. It was concluded that the first test was not run according tothe prescribed conditions, and the results were invalid and would not beconsidered further.

Figure 59 shows photographs of the primary slot piece from the tip ofthe nozzle used in the first test, alongside comparable photographs of theprimary slot piece used in the fourth test. In the latter case, the testwas twice as long and the fuel temperature was 27 K higher, thus, reinforcingthe suggestion that the rapid failure in the first test was caused by exces-sive nozzle stem temperature.

The fifth test was run with Fuel No. 2 at 491 K for 75 hours, and showedvery little performance degradation, much less than with Fuel No. 13 at thesame temperature.

A sixth test was then run for 40 hours at 478 K, a temperature betweentwo previously run tests, and the results appeared quite consistent with theothers.

A seventh and final test (No. 15) was then run for 62.5 hours at 505 K tovalidate the data previously obtained on Fuel No. 2. All of the periodic flow

105

Test No. I Test No. 1

3. 1

Iaijilc e 1

F tit-,(. I '

Port

InleI

1)0u .6

hamber

Sp-n I_

Po 1, t

I' ~(iS2 .1. I()5,I StAppea pane'ic I'uf Flc N,/ /!I II Irri . SI f(I

I M;6

calibration data acquired in these tests are listed in Appendix E. Because

of the specific design characteristics of the J79-17C nozzle, it was de-

cided to analyze three distinct aspects of its performance:

I. frimary orifice metering performance (with flow divider valve

closed).

2. Secondary orifice metering performance (with flow divider valve

full open).

3. Flow hyteresi3, caused by abnormal drag forces on the valve due

to "gum" deposits.

It was also established, somewhat arbitrarily, that "failure" would be

either a 10% reduction in primary flow, a 5% reduction in secondary flow, ora 10% increase in hysteresis. These values are considered to be high enough

to cause significant degradation in nozzle performance, without requiring

excessive test time or excessively severe test conditions.

The orifice performance data were reasonably well-ordered. However, the

valve hysteresis data were quite erratic, which should not be surprising when

considering the factors involved in gum formation and depesition. Curve-fitting techniques were applied, examples of which are shown in Figure 60.However, it was concluded that these refinements were not warranted, parti-

cularly in the case of hysteresis data. It is believed that a single incidentof high hysteresis can be just as damaging to an engine as a value obtained

from a curve. In the top curve, Figure 60, for example, a failure point of 8hours was considered just as valid as the value of 24 hours from the curve.

Therefore, the failure times of the nozzles by the three performance criteria

were established by visual inspection of the tabulated data, applying judge-ment and interpolation. In a few cases, the hysteresis data were consideredinvalid because the first value obtained was excessive. A summary of the dataobtained is shown in Table 22.

"All of the data were analyzed by a stepwise multiple linear regression program.

The correlation curves and equations are shown on Figure 61.

These fouling test data correlation equations can be utilized to estimate the

fuel-limiting life expectancy of fuel nozzles. For example, consider a fuel with abreakpoint temperature, TBp, of 533K (by JFTOT visual rating), operating at a temper-

ature, TF, of 408K. Using these data, the J79-17C fuel nozzle would be expected tolast about 7500 hours before 'failure' by hysteresis. However, applying The same

conditions to an F101 fuel nozzle (Reference 2), 'failure' by hysteresis would beexpected in only 260 hours. The difference is attributed to the smaller clearancesin the F10 flow divider valve as compared to the clearances in the J79-17C valve.

However, the J79-17C nozzle would probably 'fail' first by pzimary or secondary

orifice plugging, at about 140 hours, uder these same conditions. The F101 nozzlehas very large discharge orifices, which would not be expected to plug for a verylong time.

107

0) 4

~. 1.2

r'4 OW

1.0

r4.

0.6

w1.0

S0.94J 0 4Ij

0 0.8

of0-4J

0.6 CS _ __ _ _

to .2000 10

Cumulative Cyclic Test Hours

Figure 60. Typical Effect of Time on Performance of J79-17CFuel Nozzle with Heated Fuel (Run 3, Diesel Fuelat 469 K).

108

4) -.4 - d 0 +

4 > 43 I

0 43 .. 4~ 0 t-

C'

44 (4P CO. U 0( -.

W3 -0 -14 41~ %0 Lr -n 10 coV. C

14 CU U- r- f '0 %D C- (n r-4

0~ -A4 Z 0* A~44 U P"-4"0.,10 430 ui 0)

.3U 4J -4-4 J

to $.4 w. 430UU 040'- C

.44

0- 1--4 +

go w 00 0 - n __

1:14

43t

a)43. ; 0.4 (e'

- 004 cc C% N- N 0 -4

434

C-4 C1 e f ( i

10

100 A

- Fuel Type", ODF2

00 Je-8

Valve Hysteresis

10 Co- = 0 .2263

ca C = 0.434

y 1 00 exp0

S __________ _00 exp (x/60.5) i

0

r- 1lO~

Secondary Orifice

o CO =-225 K

Z C 1 1.226

0

E-4 1000

y 100 exp i__ __ __31)

S 0 Primary Orifice

O 0= 6.9 K0

_) CI = 0.7820

CD

100 -80 -60 -40 -20 0

Weighted Temperature Parameter, (Co+CI TBP-TF), K

Figure 61. Correlation of J79-17C Fuel Nozzle Fouling Test Results.

110=.

The equations can also be used to estimate the effects of fuel quality and fueltemperature on nozzle 'life'. For example, a IOK increase in fuel temperature wouldreduce the life expectancy of the J79-17C fuel nozzle by about 16%, as a result ofreduced primary flow, and a IOK reduction in fuel thermal stability breakpoint wouldreduce its life expectancy about 18% as a result of reduced secondary flow."

B. Engine Systems Life Predictions

1. Combustion System Life Predictions

The life analyses described in Section V-E-2 were conducted assuming aflame radiation level for Fuel I (current JP--4) and adjusting the hot gas tem-perature and film effectiveness level to achieve a match between the measuredand calculated temperatures on the panels. Similar data match curves were pre-pared for other fuels by maintaining a constant film effectiveness level and byadj-sting the flame radiation to achieve the data match. This detailedapproach permits the metal temperature to be calculated with high accuracy pro-viding the desired accurate input for the stress calculations. Although super-sonic dash conditions involved the most severe temperature conditions, only asmall portion of actual flights encounter this condition; the actual combustorlife is controlled by the sea level static condition which is therefore thebasis for the following life discussion. The maximum temperatures were measuredon the overhang at the downstream end of panel zero (see Figure 16) and data matchcurves based on best fit data (liner temperature rise versus fuel hydrogencontent) were prepared for this panel for fuels containing 14.5, 14.0, 13.0,and 12.0 weight percent hydrogen. A typical data match curve is shown inFigure 62 for simulated takeoff conditions.

The accual takeoff condition involves a higher fuel/air ratio than couldbe tested in the test rig. Fuel/air ratio trends are available in the data,including the data at cruise inlet conditions where both cruise fuel/air ratiosas well as the higher takeoff fuel/air ratios were tested. However, for con-sistency between the fuels, the fuel/air extrapolation for life calculationswas done using the methods in established design extrapolations. The resultinglife effects between fuels that result, are very close to the same as for theinitial lower fuel/air ratio data before extrapolation.

Temperature profiles were prepared for the four fuels mentioned above fortrue engine takeoff conditions. The temperature profiles were used as inputsto the SNAPTS II Computer Program and effective stress levels were calculatedfor the various fuels. The stress and temperature distributions were combinedwith available material property data (Figure 28) to predict cycles to crackinitiation. The predicted cyclic life for the inner-liner for fuels contain-ing 12.0 weight percent hyarogen is about 75 percent of the life predicted forthe fuel containing 14.5 weight percent hydrogen. This decrease in life isdue to two effects. The first and smaller effect is due to increases ineffective stress levels because of increases in temperature gradients betweenthe panel and the cooling slot. The second and more significant effect isdue to the rapid decay in material properties due to increase in temperature.The specific predicted effects of fuel hydrogen content on combustor life are:

i111

_ J

F rr

600

* Simulated TakeoffOperating Condition

500 o JP-4 Fuel Hot Surface

ICalculated0 400 Temperature.Centroid

3 0 Measured_ Temperatures

, 200

100 ... ...

0!0 1 2 3 4 5

Axial Distance, cm

Figure 62. Typical Panel Zero Temperature Distribution.

112

Fuel Hydrogen Content, Relative InnerWeight Percent Liner Life

ia. 5 1.014.0 0.9313.0 0.8312.0 0.74

The life limiting region analyzed above at the panel zero overhang doesnot show as much response to fuel effects as some other locations which, how-ever, are not life limiting regions. For example as indicated in SectionVI-A-5, the change in metal temperature with respect to hydrogen content wasonly about 8 K/% on the panel zero overhang, as compared to 20 to 30 K per per-cent hydrogen for instrumented locations further downstream, accounting forthe small predicted change in life. A slope of 25 K per percent hydrogen con-tent applied to the same absolute temperature level measured on the panel oneoverhang would have resulted in a predicted life for fuels containing 12.0weight percent hydrogen of less than half of the predicted life for fuels con-taining 14.5 weight percent hydrogen. This would have agreed with the lifedecrement predicted for the FIO liner in Reference 2. The reduced effect forthis J79-17C combustor liner is partly due to the more intense cooling mechan-isms that exist in the hottest regions.

2. Turbine Life Predictions

As discussed in Section V-E-3, flame radiation changes are not predictedto affect turbine nozzle diaphragm temperatures because of the small viewingangle. Profile or pattern factor changes would be expected to directly affectturbine temperatures, but as was anticipated, no fuel effects on combustor out-let temperature distribution were observed. The J79-17C turbine life is there-fore not expected to be affected by the fuel property changes investigated inthis program.

C. Comparison of Results

Comparisons of the data acquired in this J79-17C engine combustor studyto that previously acquired in the J79-17A and F1OI engine combustor studies(References I and 2), have been made to identify common trends and/or keydifferences in fuel effects on the three combustion systems. Key designfeatures of the three engine combustion systems are listed in Table 23, whichshows:

1. The J79-17A and J79-17C engine combustors have several major dif-ferences in design features (such as cooling technique, primary zone airflow,fuel/air mixing technique, etc.) so comparisons of these data should providean indicetion of the effects of combustor design features on the sensitivityto fuel property variations within the same envelope and operating conditions.

2. The J79-17C and FIO engine combustors have many of the same low-smoke long-life design features, so comparisons of these data should provide

113

-j

m~I 0 . --o4o

[~0 cJ~ .0

Ca01 0 0 ~ w 0 00C4- -- c1 >JJ C

~~,t m ~

to (V 4 -

.00

1.0 -3 .0 a Q

4) 0 0) 0 0 ,4)

00 0) I U~- a 0 )cu D

n)2 0)0Q 0 r_

(4~~~Q 0) OL 0 nE0 v) w,(

T0 Q) C 0 > )0 O

0.-, - 0 z~

f- 0 W ~ m .- '~0)~ <c 0) 0

0) 0)

cu0 0 =

< E-4 CCO4E

14 w'

w~ c H )oa0

'.-~ ~ 0', C O C ) 00

(114

an indication of the effects of engine design features (such as pressure ratioand turbine inlet temperature) on the sensitivity to fuel property ' ariations.

Comparisons which havv been made are shown in Tables 24 through 27. Keyfindings from these comparisons are discussed below.

Table 24 is a verbal summa.y of the major effects of operating conditionsand fuel property variations on each of the combustor parameters which weremeasured in the current and previous evaluations. The table shows that whileeach of the combustor parameter levels with the baseline JP-4 fuel are highlyengine system and operating condition dependent, fuel property variationeffects are very similar in each engine system. In particular, the followingfuel property trends are indicated:

1. Fuel hydrogen content is the key fuel property at high pressure/temperature operating conditions, with respect to smoke, NOx, liner tempera-ture and flame radiation levels and hence, combustor durability. Each of theseparameters increased with decreasing fuel hydrogen content at takeoff and dashoperating cond :tions.

2. Fuel aromatic structure, as indicated by fuel naphthalene content isa relatively less important fuel property. However, at high engine pressure/temperature conditions, some effects of fuel naphthalene content on flameradiation (J79-17A only) and smoke emission (79-17C only) were observed,when the concentrations were high (12 - 25 volume percent). Peak liner tem-perature levels were virtually independent of naphthalene content in all cases.

3. Relative spray droplet size (calculated from fuel viscosity, densityand surface tension) is a key fuel parameter at low pressure/temperature opera-ting conditions. Low temperature relight capability (both cold day groundstart and low Mach number altitude relight) decreased with this parameter.Idle and cruise CO and HC emission levels also increased with this parameter,alone in some cases, and jointly with decreasing fuel hydrogen content inothers. In the one case where high temperature atmospheric pressure testswere conducted (FO), pattern factor increased with relative spray dropletsize. However, in full density tests of the J79 combustors no fuel effectswere found, and none are expected for. the FI01 at full density.

4. Fuel vaporization characteristics (as indicated by boiling range)of the fuels tested are highly confounded with the atomization characteristics,so some of the combustor parameters correlated by relative spray droplet sizecould be alternately attributed, at least partially, to vaporizationproperties.

5. Fuel breakpoint temperature (by JFTOT) is a key parameter with res-pect to fuel nozzle fouling, and hence, durability.

Table 25 presents a quantitative comparison of the effects of fuel prop-erty variations on exhaust emission characteristics. Coefficients (ko, kl,and k2 ) determined by regression analyses are listed for each engine, operating

115

V

II .1

COy~ . I-

8 t

*~ ~116

* N 00000--Oo 00000

.14 0000 000004 00000

I 00000 00000 00000

a ~ l I0 N.... l O

44

oo -; C;

1.,.1 40' I4 N'C" -4 W

Li i V, 1111- 1 14

01

0

0 141117

cl, ~ ~ ~ .r v- 00m n -- x )1

'4.4 - 00000 0000-. 0.00000

.) . . . .....

I. 00 0 00 0 00,000 041 __ _ __T_ _ 00(0 a, C1

!L+ C1 iCii C )C.I .. . ....... c

co4

W &4 S

00

C4-

C14~

Aj + CIA '4 % ' m 0 -8 044

0) (

) 0.

U)~V (vO~. t0'O-*

. N 0000 0000 ai -- w-4.4 4 .1 0 04 0~

0 0000 00000 -(o ++ +++ 4-..'

00 I 4 Na) a0)

C.)~~J C' . 'J 0 04.4 0'' '1' C 1: ca 3- a

U..4 ... M .. C' ) 1 o0%0041 1 111 C 0 a 0

U) C4

00 w w

00 00 34 wa'o~ I 0 S.c II o*

1- ~ 4 '4C44* m .CII 0) 44 CC '

.4- 0. Q) 0) (0 ) n 0

m. >- m 2 4c

0 Q) IVCCC U 3.. CCO X 0 Co

0 C. mE 0) f.- Q) 0)4 0 0(0 .

0 2 2

118 ., '

-- Aou

condition and type of emission. In each case, the first constant (k.) re-presents the emission level with current JP-4 fuel; the second constant (kI)represents the sensitivity to the primary fuel property (generally hydrogencontent); and the third constant (k2 ) represents the sensitivity to any otherproperty (relative spray droplet size or naphthalene content).

Comparison of the coefficients in Table 25 shows that, as noted previously,the emission levels (ko ) are highly engine system and operating condition de-pendent. However, the fuel effects (kI and k2 ) are relatively consistent bothin their presence or absence in each case. The magnitudes of the fuel effects (kIand k2 ) are also fairly consistent, and the differences are probably related tobasic engine combustor design features, such as pressure ratio, stoichiometryand residence time.

Table 26 presents a quantitative comparison of the evfects of fuel prop-erty veociations on parameters which influence combustor 1,ner durability.Again the listed coefficients were determined by regression analyses and havethe same representation as those in the emission comparisons. In this table,however, the fuel coefficients (kI and k2 ) always represent the linear slopeof the normalized durability parameter [qr/ko or (TLM - T3)/ko1 to a decreasein hydrogen content (kl), or an increase in naphthalene content (k2 ). In everycase, k2 is at least an order of magnitude less than kl, and generally theydiffer by two orders of magnitude. These comparisons quantitatively illustratethe general insensitivity of these combustion systems to fuel naphthalene content,at least within the ranges tested.

Table 27 presents a comparison of the effects of fuel hydrogen content onpredicted relative combustor life. The relative lives of the J79-17A and FI01ijcombustors are expected to be quite similar and sensitive to fuel hydrogencontent, even though they incorporate vastly different cooling design tech-nologies and hence, much different absolute lives. The relative life of theJ79-17C combustor is predicted to be significantly less sensitive to fuel hy-drogen content which, as described in Section VI-B-l, is attributed to thehigh front end cooling.

1I

I) 119

-" - . . . ..- . . .. . . . ... . . . . . .. . .. . . ..-

Table 27. Comparison of Hydrogen Content Effects on Predicted

Combustor Liner Life.

Relative Combustor Life

Fuel Hydrogen Content J79-17A J79-17C FIOI

Weight Percent (Reference 1) (This Study) (Reference 2)

14.5 (Current JP-4) 1.00 1.00 1.00

14.0 (Current JP-8) 0.78 0.93 0.72

13.0 (ERBS Fuel, (DF2 0.52 0.83 0.52

Reference 5)

12.0 (Minimum, percent 0.35 0.74 0.47

these Programs)

120

SECTION VII

CONCLUSION AND RECOMMENDATIONS

Based on the J79-17C combustion system experiments and analyses conductedin this program together with comparisons to previously conducted J79-17A andF101 experiments and analyses, the following conclusions and recommendationsare made:

A. Conclusions

1. Fuel hydrogen content strongly affects smoke emissions, liner tem-perature, flame radiation and NOx emissions. Hydrogen content is,therefore, probably the most significant fuel property, particularlywith respect to high pressure/temperature performance, emissioncharacteristics, and combustor liner durability (life).

2. Fuel Vaporization and atomization properties become more signif-icant at lower temperatnre/pre.asure conditions. Cold day startingand altitude relight capability are highly dependent upon these pro-perties. At least for the fuels tested, these properties do notcorrelate with combustion results as well as the relative spraydroplet size calculated from fuel viscosity, density, and surfacetension.

3. Within the range tested, fuel aromatic type (predominantly mono-cyclic xylenes or dicyclic naphthalenes) had relatively littleeffect on combustion characteristics.

4. Fuel breakpoint temperature (by JFTOT) is probably the mostsignificant fuel property with respect to fuel nozzle fouling.In accelerated cyclic tests, fuel nozzle fouling rates weresatisfactorily correlated by this fuel property.

5. Fuel property effects on combastion system performance and durabil-ity generally tend to be similar for the J79-17A, J79-17C and FI01engines, even though the design features, operating conditions andabsolute performance and durability levels differ significantly.The most apparent exception is in respect to the decreased sensitiv-ity of the J79-17C combustor life to fuel hydrogen content.

B. Recommendations

I. Selected engine tests are recommended to verify the trend estab-lished in these rig tests and analyses, particularly with respectto liner temperature, flame radiation, and hence, durability.

2. The fuel nozzle fouling testing in this program appears to havebeen successful in establishing accelerated test techniques and

121

data analysis methods. Additional tests with greater variationsin fuel and air temperatures, ftvel types and longer cyclic testsare recommended to validate the procedures and results.

122

APPENDIX A

FUEL PROPERTY DATA

A. Detailed Physical and Chemical Properties

Thirteen test fuels were sipplied by AFWAL/POSF for evaluation in thisprogram. Detailed physical and chemical properties of these fuels were de-termined from AFWAL/POSF retained samples and are listed in Tables A-1 andA-2. Selected data from this table together with a general description andrationale for selection of the fuels utilized in these evaluations are pre-sented in Section III of this report.

It should be noted that as before (References 1 and 2) total aromaticsdetermined by fluorescent indicator adsorption (ASTM Method D1319) were alwayssignificantly higher than the values obtained by spectrometry (ASTM MethodD2789). ASTM Method D1319 is not recommended for fuels having end distillationtemperatures over 587 K, but only the diesel fuel exceeded this value. ASTM

Method D2789 is not recommended for fuels having 95% distillation temperaturesover 484 K and all of the fuels exceeded this value. Recognizing this fact,Monsanto applied an alternative method (their No. 21-PQ-38-63) to five of thefuels shown in Table A-2. With the exception of a few of the components ofFuels 8A and 9A, the results by the two methods were not substantially dif-ferent.

Samples returned to POSF were checked for any contamination that may haveoccurred through shipment, storage or handling, by glass capillary gas chro-matography with each returned sample's retained counterpart. Some returnedsamples showed variation from their retained samples, but in all cases thesedifferences were not significant, inasmuch as they did not adversely affectany test or related data analyses. Major differences were noted in returnedsamples of fuel 2A1, dated 15 Jan 80 and 8 Feb 80, from the Fuel Nozzle Foul-ing Test Rig. These samples, which differed in high boiling components, C15and above, also had lower JFTOT breakpoint temperatures. The apparent causeand effect of these differences are explained below.

B. Computed Combustion Parameters

Table A-3 shows several fuel parameters which were computed from thephysical and chemical properties for use in conducting the combustion testsand analyses of the results.

Fuel hydrogen-carbon atom ratio (n) was used in the exhaust gas samplecalculation. It was calculated directly from the hydrogen weight percent(H) by the relationship:

11.915h (2)n = 100 - H

123

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I

F 1I "-'-C COO -Co

'.00 - 0 ;0 ~ I 0 '-CC 440''- C~ o-o-

0-j "' ::::x:

000 -- IC 440.l00 404440 I *4C~C~ *S..~

-' CC 9~,~'~ie.-o ... 4404 C S - 000 Ceo- ~

004 IC C 0 04 00010 0,04.014,- CO 00 4 c400 COt '-"44,. - I -C -- 00 04C

4 004000' CC CO CC 0.. 0C0~0 c-oo ""C

4'-40 C C.t 00041 COO' ... CCIE~ 10 _______o.'. C4~OC C

-- 00 44O~~CI0 CC 001- CC

S ICC C - CnO0 c C C... 00~ C CO COcOOlc... ,

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0.4J *44-0 ~ 5~, * 0 - C'C~ *44.0 c00 CCC..c 440 00

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0 ________ -- ____ -0.0 occ. - C 44 04 O~. 0 - OOOOC

3.4 C 40', '0 0 0 0... 0044 4..004 c Oct04. CC-C 000.

04 00 I 400 -- C044O.-,0044444440-~..0 0000 00CC ae 0

CCs-- ~C .0 S0 ocno-4 404 C- o0"Oo e - - - s-Cc 0CC'-0 0.......4 014 oOCOCC Cc4cO I Ct-C CC'-4 01C... '0 4 00 . 0n4i 000440...'O iC'-c -C~00 . -

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0 41 ~.0C -- 00 0 440*00 440444 -- ~~Oo 000 0000 0.04

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C -________ CC0 CC 4... o , ~ 00100 C

0 0004040,0 04 - - CO Co c.-.c.-in Coo OK

.~04t IC 0 e 40 .. 0I004o50 .- Ot-.CO . 40.- 0.044 -C* en -- t -- C4400-.InI0l0 4 4400 00 000 '-""'-4. 40

- .-- ___________ ______ _______________

I .. : j ,..5 5~ ~j5j I ~aC U 00 0 U gCOU~ ~ 00000 0000 --

.- 2~ t~0 30 --- _________ .- 2,-a

r-4 Cc

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a.CCCOC C 9if-4 - - - - z CCCC Ca-a - ax '~14C CC. - I C -- C .~

CC4C.-CC10b.4OtC~ * - IaCOaa. - C CC *4IC C-

CC..aC CCCC> C 0 CCCCC - CCCCCa..CCt00Ccsaae0aCC~CCC.IO CCC IZO...CC CC~ o a. ~CC'-C'-CCC - OCCOOC-

IC *..tCI)C4 1' 3a.Ca.C ~ ~ 4CC xOCCCC o00n.~0oo.--0o-.o00.I CO -COt-- C:.** CC .40CC C ... O4CIC tt

-- C .. Ca*CCCa C C C 3~'a n :. a a a

124

NIT p.-.

O0, ~ 0

.. a-0

v

C44

a UN* G0NuN

V ~ 0

o .to

A ~ 10NN- 0!4

41

0) )

*0 0

*0 -0

co

00 C 0'C

m~ eq

~ 0 Q~O ' 0 to C:

.V0

4~-. Nr

1025

CC

Mt It54 1

40.

0 r - ,( 0 U) M W t0 ('0 ON '0 I

-~0 0 0 -4 11 (V) -, N" 0 0 0 t 0 '

i) 00 nC1 1 nC

14 4,

;4 0 , $.1

a a 0 00 Co r- 10 4 1~ O u 0 0 4

'4. g 4) a 0C - y

0a W

-. 4

4.) '-04 0) r 1 t 0 n m a

u ,4

-4-0 0 00 . 0

00

co .0 -

4.)000

0 ________W A

C.) uF 0

.,126

and ranged from 1.616 to 2.017 as hydrogen content increases.

Stoichiometric fuel/air ratio (fst) was used to calculate comparativeadiabatic flame temperature. It was calculated from the fuel hydrogen-to-carbon ratio () by the relationship:

f 0.0072324 (1.008 n + 12.01) (3)st = (I + 0.25 n)

which assumes that the fuel is CHn, that the air is 20.9495 volume-percentoxygen, and that the air has a molecular weight of 28.9666. For the testfuels the stoichiometric fuel/air ratio ranged from 67.52 to 70.31 g-fuel/kg-air as hydrogen content decreased.

Stoichiometric flame temperature was used in analyses of NOx emissions.It was calculated at takeoff operating conditions (T3 = 664 K, P3 = 1.359MPa) using a standard equilibrium-thermodynamics computer program (Reference14) and ranged from 2494 to 2516 K as hydrogen content decreased.

Relative require.d Euel flow rate was used in all combination tests toadjust the JP-4 fuele engine cycle operating fuel flow rates for the reducedheating values of the other fuels. The factor is merely the ratio (Qjp_4/Q)and ranged from 1.00 to 1.0375.

Relative fuel spray droplet size was used in analyses of the low-poweremissions and relight performance. The J79-17C combustion system employshybrid pressure-airblist atomizing fuel nozzles, but even at relight fuelflow rates, the bulk of the fuel is injected througn the secondary airblastatomizers. Therefore, Ei-Shanawany and Lefebvre's correlation parameter forpure air atomizing nozzles (Reference 15) was used to estimate the relativefuel spray droplet Sauter Mean Diameter (SMD) from the test density (p), sur-face tension (a), and viscosity (v) by the relationship:

SMDjp- 4 uJP) ( 0. 1 vJD_4 PJP-4 0.4

I + C2 0) .

aJP-4

where: ( 0.015 ~ 0.1 1.2 0.7C2 0 7 J\Dp UA PA /220.0737

127

and Dp, UJA PA are atomizer diameter, air velocity and air density and allparameters are expressed in basic kg-m-s units. For these relative spray

droplet diameter comparisons, J79 idle operating conditions were assumedfor which C 2 = 381.8.

As shown in Table A-2, none of the blending agents appreciably changed

the predicted relative droplet size of the base fuel. However, the JP-8 based

fuels are predicted to produce mean droplet sizes about 38% larger than those

of the JP-4 fuel. Further, the diesel fuel is expected to produce mean drop-let sizes about 66% larger than those of the JP-4 fuel.

C. Thermal Stability Characteristics

When performing hot fuel tests in nozzles (or other fuel system com-

ponents) it is essential to know the actual thermal stability of the testfuels. This is not done in routine fuel analyses, since the fuel specifica-

tions require only that the fuels pass the thermal stability requirement at

533 K when using ASTM procedure D3241 (JFTOT). In order to determine theactual thermal stability, it is necessary to run additional tests at higher

temperatures until a temperature is reached at which the fuel forms heater

tube deposits of Code 3 or darker, and/or a filter pressure drop of 25 mmHg or more in less than 150 minutes. The highest temperature that can be run

without reaching these values is known as the thermal stability "breakpoint"temperature. It is not uncommon for aviation fuels to have breakpoints that

exceed the specification value by 50 to 75 K.

When selecting fuels for studying the effect of fuel thermal stability on

component performance, it is desirable to select two or more fuels that dif-

fer significantly in breakpoint temperatures. The precision of the JFTOT pro-

cedure has not been established, but experience indicates that repeatability

of the breakpoint determination is probably no better than 5 to 10 K. There-

fore, fuels differing in breakpoint by at least 20 K are desired.

It is also desirable to show thermal stability effects in a reasonableperiod of time, without applying unrealistically high temperatures. This

requires using at least one fuel with a marginally acceptable, or even a

failing thermal stability. Such fuels are extremely difficult to find.

It was originally proposed to use a JP-4 from a refinery which has

occasionally produced low thermal stability fuel by a copper-sweeteningprocess. However, when a sample was checked in June, 1979, it was foundto have a breakpoint of 573 K (40 K above the specification requirement).

Another candidate, low thermal stability fuel was located at a west

coast refinery newly engaged in jet fuel production. This one has a break-point of about 511 K, and was of sufficient interest that the Air Force

planned to secure and store a quantity for future testing. However, before

thie could be accomplished, the refiner changed his processing and improvedthe fuel quality substantially.

128

No It was ultimately agreed that a satisfactory low-quality fuel would be aNo. 2 diesel available locally. This was estimated to have a breakpoint ofapproximately 513 K. The higher quality fuel proposed for comparison was aJP-8 with an estimated breakpoint of approximato y 598 K.

Emphasis was placed on dete-mining the breakpoints of the twofuels used in the nozzle fouling tests. When it was determined that thebreakpoint temperatures of the returned samples of fuel 2AI were about25 degrees less than their retained counterpart, samples of fuel 2A1

were sent to two outside laboratories, Alcor and Exxon. Their results,as well as POSF data, are shown in Table A-4. Acknowledging the differencesin the returned fuel's breakpoint, it was decided to use the returned samples'breakpoint temperatures in any of the required data analyses. It was laterascertained that another JP-8 fuel, 2B, was used instead of the suppliedfuel 2A which was, at that time, nearly exhausted. The reported breakpointtemperature of fuel 2B is 573 Kelvin which matches the returned samples offuel 2A1. Later, a second quantity of JP-8 was requested from POSF fordelivery in early April 1980. This fuel, drawn from the 2A1 suurcc tank, wasdesignated as fuel 2A2. Returned samples of this fuel have the same break-point, within the accepted limits of repeatability, as its retained sample.Fuel 13A1 breakpoint temperatures also match their retained counterparts,within repeatability limits. Fuel 13A2 represents a second quantity of fuel13A which came from a different source than 13A1 and was not expected to havethe same breakpoint temperature.

It should be noted that, in every case, breakpoints were based on pre-

heater tube deposits since filter pressure drop did not reach the failurelevel until after the tube deposit reached a visual rating of 3.

129

Table A-4. Thermal Stability Rating of Test Fuels.(ASTM Method D3241)

Fuel JFTOT Breakpoint(l), KNumber

Retained Samples Returned

SamplesUSAF ALCOR EXXON USAF

1A 5382A1 593 653(2),553 603 573,5732A2 - - 603,603

583,5833A 593 - -

4A 578,563 - -

5A 583 - -

6A 573 - -

7A 560 - -

8A 538 - -

9A 541,533 - - -

10A 543 - - -

IIA 576 - - -

12A 553 - - -

13AI 533 523,523 523 523,52313A2 - 503,503

(l)Defined as the highest temperature at which a maximum visual

rating of the heater tube deposits is less than a code 3.

(2)Sample contaminated by red gasket material.

130

201

16 Correspondin~g ATDR ______ __________

12 0~

00-o _00

00

00

0. 0-

4-

00

-W Breakpoint-H

.0

[-4

0 (3D 0 Q& -CD-

40520 560 600 640

Tube Temperature, K

Figure A-1. JFTOT Results on Retained Sample of Fuel No. 2A1, by

USAF. 131

APPENDIX B

HIGH PRESSURE TEST DATA

Table B-1 presents a summary of the key reduced data from the high pres-sure performance and emission tests. Table B-2 presents additional measuredand calculated parameters for these tests.

Table B-3 presents the measured and corrected CO emission test data atthe four engine operating conditions. The procedure for ccrrecting themeasured data to engine conditions by the ratio of the severity parameters isalso shown. As can be seen, the corrections are all relatively small except

at dash operating conditions were the test pressure was reduced. Measured andcorrected hydrocarbon emission data are similarly presented in Table B-4. NOxdata are similarly presented in Table B-5. Small differences in measured andcorrected NOx emission indices are primarily due to the humidity correction.

Table B-6 presents the smoke data analysis. As described in SectionVI-A3, no operating condition severity parameter could be readily found, sothe data are presented as they have been corrected to engine Station 8 fuel/

air ratio conditions, described in Appendix F, and averaged at each simulatedengine power level.

Detailed liner temperature measurements are listed in Table B-7 (innerliner) and Table B-8 (rear liner) for the high pressure performance tests.The data are presented as liner temperature rise (Tliner - T3). The indicatedthermocouple locations correspond to those shown in Figure 16 and Table 7. Cor-relations of the liner temperature rise data with operating conditions andfuel hydrogen content are summarized in Tables B-9 and B-IO.

Table B-11 presents a summaty of the flame radiation data analysis.

Linear regression curve fits of the data (see Figure 50) were made and fromthese the quoted engine flame radiation levels were calculated.

Detailed combustor exit temperature profile and pattern factor dataare listed in Table B-12. Linear regression curve fits of these data (seeFigure 53) were made and are summarized in Table B-I, and are the quotedengine pattern factor levels in Figure 54.

132

Table B-i. Basic High Pressure Test Data.

Av1 79 $3 0.5 ± .1 S3 .

4 2.s 2.2 :6 50Is 0. : 11 5 b 6b 1.

I '-, I 0 -' I

0 10 .. 379- t, 4 . 95 i t.! .I -'13 119 '

,11 6.16 1.7 656 ',. t 172 9 99

14 51 11to 794 1.535 1 .1936 2 9,0 0. IA 3.31 ',s 389 102 o 1 . .". 'o,, .4 s4 11.2 99 13 '9 .2 .8 .1 .3 43 75 12 32 95 93 9 1.1%5 0 i?4 91 23 4 0 4 2 3 20 0 1 4 36 3 1

6: 25340 2.327 0.195 380 13.8 1.0 3 .93 - 310. .3 4 4 69 5.2 1IA 91 $7 9 5 . 464 9301 7 12 0037 446 15.8 0.6 0,3305 13 0 0 13.' 2 3 3 95 .5 16

6 .3 .6 .32 85 3 ..3 13 300 0 120 I0 341 419 1 0 14 333 3 9 4 3

8 4S27 ::.340 1.9 77 33. 1 . 3 91 t 406 1099.0 1 3.330 0.139 303 0 0 22 1 1 7.2 99 5 .S .30 2.457 0.197 $70 13. 1, .93 :5 498 99.2 1. 224 0 3341 9 19 0 0 o 3 6 36 1. 99.6 2

14 1156.7 3.634 1:161 769 2 12. 1. 6.0 1 111 M9. 1.360 0 33 30IS 1: 0 0 233 14. 99.9 .

33 1 .3 6023 1 .6 2 49 43 36 13 33 7 0. .1' 0.534 1.4 0 9 % 3 3. 9 7 3.

14 3 .56 0 .4 0.269 343 79. 2 1 4 3.2 58 3. 1,1840 0.29 35 1. 4 3. 3 8 19.8 2 .3 1

16 2 .419 1.8 .6 3 . 1.1 3.2 1 - 21 2 1.8 .2 26 : 60 .4 39 23 32

6A1. 3 2 1.821 3.:7 0 .3 47 7. 0.4 . 8 38 92. 3. 0.333 197 100. 45.$ 26T 92 9)374 1.I39 I I .1 1. * 0.3 418 9.9 .3 . 64 1: 383 9IT1.0'.34 7 9. o3. 2. I, L 3. 93 3 '

1923.0 .4116 0.67 360 14.0 '4:7 3.87 S 16 Is. 3 .097 0 2803 7 : 32 4 30. 01 )2. 16.b 096 1 3 .

20 1 2938 2.312 0.874 33 90 07 74 9 9.9 3.4 0 : 2103.. . 0. - 23 4 9 9 a7 27.7u~1 61.3 .0'334 87 3.0 . 7 33 32 30. 110 031 40 2. 0.2 16.0 373 999 14.3

22 7 7.43 .9 17S 622 13.5 0.7 3.32 S 86 103.1 t.9 0. 482 3. 0.) 23.8 19.4 "9.94 33 223 8 8.338 3. 69 1.198 775 12.9 0.7 3.1 471 102.11 1.108 0.311 3'l 1 : . 4 80 599 .

)19634 .7 1.200 776 14.9 0.3 3.84 $ 34 101.7 1.106 0.3316 '.0 0.1 0 .2 2.3 t 1 3 596 11.3

3A 23 9 6.S3 .0 .0 748 15.9 0.6 399 $574 302 .0 0.3 3 09 0 - 7.3) 9.98 4-.26 0 677 141 1.198 77 t27 06 39 4 10 .31: A.4 430 0.9 0 26 3 59

277 4 70 4 83 1.273 48 13.1 0.8 3.2 33 34 .0 .0 3 1.0 49 1.3 0 3. o7 3 99.9 12.

20 7 7441 .307 %.345 846 1. 0.0 3.39 1 7 302. 1 :1 0 1 40" t2.7 0 3. 1 7.3 9 97 30.4

29 4 5 4 2. 2 .465 7693 0 ~ 0.1 3.53 4183 94.2 1.14 0.24 311 2 0. .7 3. 99.0 7b .30 2.4 .13 0.72 314 34.3 0.5 3.3 U 88 9.2 1.28 0.3314 9 . . 40 9 0 3.

31 2 1.626 1.S73 0.252 623 5 0 0. 4.4 - 27 6. 120 034 39 923 2 4 5 94.0 332 3 .87 .6 .23 47 9. 0. 3.7 144 90.4 3.212 0.442 132 -0.7 38. 2.2 1 94.31 6 .

2IA14 3 8.65 .638 t.202 7503 12.7 0 7 3.96 - 54 101. 3.103 031 494 0.9 0 " 340 99.95 .

3567 7.49? 6.32 1.8 63 1.9 01 31 06 10. 3.0 0.8 40 3. 0. 166934 743 5265 1.34 61 13. 0. 1 3.6 - 17 104,2 1.137 0.287 404 1.3 0.1 . 301 5.0 230 6.7 3.0 119 70 1,4 .7 .4 32 116 109 0.2 43 300 99 2 .0 96 9 16. 0

37 4 6.3 .2 .47 36 .0 .3 3.6 - 5165.13 033 99 9.4 0.2 633 14. 9.7 36

Is53 2.133 2.482 0.47 33 20.6 0.1 3.49 - 720 6797. 3.2 0.4 293 5.6 0 2 9 . 32'9 939.719 33 6

39 2 3,30 1.313 0.233 45 4.2 0.6.0 8 6.9 1 .3 0 390 374 3t0 7 70 6 ,j, 32 91 9 4 3140 2 1.728 1.53 0.233' 43 , 31 1 .5 - 30 407 210 023 6 17 3, . 32

42 1 5.800 1.523 0.252 42 0.,. .3 33 23. .0 .2 7 7b.2 24. 2 . 0 93.2 043 4 701 2463 0 47 36 13. .3 0 5 9 36 1.6 3.12 09 0391 237 7. 0 04 3. 93 3

IX 44 3 2 . 79 0.27 37 1 4 4. 3.3 - 67 0 .3 107 0.33 633 01. 0 .4 303 9.44t I 7.:o 50 .2 3 0 12 683 33.0 4 9 S 2.9 1'71 10003 019 0.73 .0 2.2 0. 137 7 9 9 3

47 0 .389 4766 1.17 77 01.3 1 .0 . 47 218 3.1 0.0 7 0 0 0 22 0o 33 3. 6. 99 95 3.9

43 4 S 2 .'9 0 33 0.238 435 10.0 4.9 .4 1 33 24 1.333 0.293 101 99.4 49 24 2. 740.30 2 1.1 .34 0 27 4 15 8.0 4.9 11 274 33 31 0309 3 030 9.2 0 11,.3 .7 9...20 40

7346 293 243 047 36 3. 4 . 3.8 2 05163 10. 1 0 0.22366 34 0 1 2 Is 3I13 9 7 7

56 3 9.50 3.36 1.85 776 12.6 3. 04_j 4 0 1 3 09 30 4 0 9 2 0 1 3 $. 233 49 5 9

13 ~~13 111

Table B-i. Basic High Pressure Test Data (Concluded).

-9 :M 0 .2 1 .6 942 4.06 408 2 7a It33 is,* I1 b.e 1.: $.0 6 3.2 12 0. .1 0.2 1.4 6.5 26. 465.0F00 2 2.620 1.0 0.1 01 . 30 ?4, 27 90. 4207 0 ' 2, 2 2 .9 9q 30 24 1560 0 27 2 1 (.2 *3 1". 2. 1 ... .. 02 0.12 o2 22 00 01. 0 93 200 99* 2

02 0 ;.057 *027 1" 19 10 1 1'922 2 22.9 232 209. 19 0 29 9 . . 02 .4 99 it I702 7 5 t00 I2. . 7 8 1122.8 020 t00.0 :.09 0.0 l 22 'l 01 0 30 2. 9 92 22 0 I.

*3 8 3020{2. 29 9 222.3 049 201.5 1 024 0.223 076 2.1 0. 25b202 9.9 .*0 9 6.430 20 23.2 1. 3l 5 Il . 3 IM 3.3 490 202.3 2 02* 0.208 095 t.0 1 0.0 23.3 10.0 ~9 0 . I2.

II s : 6 I$ s I* ~ 10 I.5 3 43 .22.0 442 99A . 011.8 0.2 M 0 0.9 9. 10 .

.2229 05 2.1 1.2.0 1 3920.I

. 2 39 0.219 032 9. . 2.3 3 11 : 8 3. 0 1.2 1.oat 0.2*21 41 62.0 302 . 1.3 19:.1 0 222 2 2.033 2. 1 1 0.00 1 01 2. .2 .8 280 11 2.12 0.0 29 93 42 21 . 32 .

I 3 3 2 225 2. 4 0 . 3 020 9.9.7 2 2 8 2 4 8 .0 1.092 0 .212 209 8 0. 4. 2. 20 2 9. 20 3.2

23 23 240 0.4.77 000..22.0 .7 3 0.04 83.7 7 0 91 90 0 .*0 370 22 1 .2 2* 9. 20.2 99.42 4.2193 29 3 2 0.48*6 020 12.8 1.3 2 309 8.. 0* 27 229 0 4 00 82 0 .0 to2.2a 99.20 2.0

8l0 1 .33 .493 2.2sa2 3 t . .1 $. 270 20.8 7.1.0 - 9 9 . 27 0. 21 109 01.9 0.2 24, 1 S.6 $37 999 2.0

89A 8 6 *3 3.37' 2.383 7U0 23. 3. .9 1 3 029 202..* 3 2 . .0. 23.0 oI 4. 19 9 7

I. 1 " ,.4 0.3 .29? 7 * 264 .4 2.2 30 2. W 0 047 Lo 06 0.26430 07.5. 09 9 0 4.

89 1 .10 2.6 80* 30 . . 3.7 90. 12 98 0 98 0.222 20 12.7 0. 0. 29. 9.7 27a87 3 .790 22 0 29 43 . 2. 3.9 8. 304 03.2 t 0040 0227 2 2 4 3 . 2 8. . 24 2. 08 .

is 2 .503 2.444 0.39 J 00 2S. 2. 1 9 82 30 to" 0190 024 1299 7. 30.2 4 2. 9 92910 31.0it 89 2 .803 2.09 0 I 3 422. 20 40 8 3 0: .0 0.2 2 230 202. 09 21 . 22. 2so90 .2 4 2474 03 02 20.e I.0 2I . 2 82.6 373 9 . 07 0 230 203 202. 30. . 0 22 946t 90 0 2.

920&A7 40 0427 let 22. 2. " 2.005 90. 1 2 0 2103. .8 8 6 2.2 0. o 31: 29.2 99 94! 8.

: * 9 . 9 3 .4 0 2.2 2 7 0 2 2 . 1 . 3 .0 s 2 2 9. 2 I t s 2 0 3.2 2 0 4 . 00 70 2 .0 0. 2 3. 2 8. 2 9 4 2 .

190 '23 020 2 0 *2 2 1.2 2 27 222 032 20* 7 .: 04 9 , Il2 24 02:22 2 23 27.9 99.90 22.320 7I720 0: 22 LA38 99M93 22 .0 3 go.) 200 6 2 0 0 0.210 02 0, 20 0 20. 2 9 99 27 2

202 ) 2.82 00 7 * 0li 2 3.7 2. 3 72120 30 992 220 2 6 387 .7 2.7 2.4 2. 1 99 82 222Is2 2 l0 3.62 0 7' 30.25 42 0: I 30 .8 2 1 702 98. 2 07 0.200b '23 7.3 0.21 927 23.0 IS 02 2.

203 3 2.64 2.002 0.2 9 026 9 0 .0 3 :1 8 2 962 90 . 9 2 01 0.2 10 2 0 2 2 0 0 ' 0 2 3 9 .:1 11 11 "1 4 .0t29 ,.1 .. I : 0 607 273 0426 p 221 0 - 34 143 02 99 I% 31 912 79111: 1 11 l.20 .0 9 022 2 08 '242 93 2 2.0 326 96. 222 6 402 -200.0 0 0.20 43 0.2 012.0 2.2 99 4 26

I0'll 2*8 1.00 1:2 '9 3. 2.2 3.7 2 13 l . 1-000 0 9 0.2$0 57 0.0 0. 201.0 2 2 99. 2.):3 9 7 37 0.30 .39 *9 2. 22 30 0 31 0 -: 224 022 0222 42 21 0 .) .2 1 7 2 4 9 9 99 28 920 7 0 j 4.*82 .29 4 2 209 .3 .0 1 10 2 0. to 0 9 6 0.2 0 1 2. 0:2 23.0 29.0 I 99 Z 1

0. 1 2 O 88 0 73 080 22.32 .1: 3.20 920.2 090 20081 .092 0,229 209 2 0 0.0 0. 23.2 9 I C$

7A 0 1 6.U 2 9 208 0 28 39 922 . 3 57 122.2 9 7 20s05.9 077 0 207 2 7 2 02 90 2. 9 2 2.222 1':026 249 022 03 . . 1 4.4 8. 3 02 220 0 33 20 0.0 400 1 9" 20'3No

02 .1, 152 7 0 2 .21 3 7 2 3.6 262 9.0 2 2 0 039 I 0 20. 29. 2 8 9 92 22 2 22.

,134

Table B-2. Supplementary High Pressure Test Data.

f I..

oI A0 IWO wx z

2A 1 2. .0 28 1. .2 8 0 .1"420 .018 0.9 0.19 .2

4 85 1.1 29 91 1.2i1.9 .4 8. 0.054 013 0.43 0.95

1. 2 1 1.2 1. 8 0.3 1 1o5

2 28 00 . 1 a 0 0 . .0-' 0. 472 1 7 1.; 0 1 0a 00 18 .0

01 - 1 0 0N ,. 00 14 .

40 3 ~ 0 41 ~ ~ .~ .~14.. 1.0

.3 2 1. 6 001.. . 2 1 .0

U. I A. o

1 2.9 1.3 29 14. 8 0.90 0.38 0..110 20 0 1 O

6A 7 2 .91 .1 28 1 .2 0. 28 ! 0 328 1 .40 7 , 0 0. 7 0:976 0.192 1.2125 -

8 23.6 1.27 289 12.0 0.862 0.319 0.20 4 1 .35 09 0.153 I 103 27 1.2 288 34.8 00 010 it 0..15 I 1.

41 34.5 124 28. . 0 1.34 1 0.1 6 0 0.07 0.019 1 L43 1.0 8

529 0.1 3 1 001 00 22 81 1064

28 33.9 1.30 288 14.8 1020 1.680 0.293 2.3 0.018 0.01 1.72 1.01 -1 25 33.6 1.27 289 12.0 1.8 19.773 0.277 1 7 0.039 1.049 0.783 .060 -

30 38. 1.22 288 7.8 1.158 0.59 0.610 8.4 0.014 0.012 1.02 1.034 -17 28.5 1.15 289 86.1 1.52 1.072 0.624 8.7 0.03 0.019 1.23 t.07

12 29. 1.3 27 8. .25.2 .3 03 001 001 .0 .5

28.5 1.1 1 28 6 100.8 1.669 1.9 7 1.023 1 60. 0.05 O.01 79 1.204 1.0423 2 6 1.71 288 3.9 1.35 10.5 0.12 3.1 0.091 0.122 0.5.5 1.061 2.9 1.16 288 46.9 1. 28 1.642 0.29 4.3 0.022 0.0 120 0 4 1.033 -IA 39 33.9 1.14 288 70.9 1.3 1.542 0.285 3.3 0.017 0.012 1.784 o.056

10 32.6 1.14 288 7.30.7.4 0.36 0.10 3.0 0.026 0.012 1.7 36 1.0351 28.9 1.15 287 86.1 0 1. 1.7742 69 8.80 0.038 0.0219 1.242 1.07112 28.6 1.15 287 98.5 1.663 1.967 0.848 10.5 0.058 0.018 1.205 1.01313 27.6 1.11 288 33.9 1.081 0.830 0.136 1.8 0.016 0.157 0.505 1.07414 27.2 1.12 287 46.9 1.22 1.082 0.08 1. 0.10 0.12 0.463 0.993021 25.6 1.16 289 51.9 1. 1.0.6 1 0.081 0. 0.010 131 1.340 1097716 25.6 1.26 28 11.3 0.796 0.69 0.16.0 2.3 0.021 0.03 0.120 1.186 8

A 1 23.9 1.2 288 12.6 0.828 0.328 0.39 3.0 0.01 1.101 0.14 1.124 24.0 1.17 286 14. 0.851 0.0 0.63 4.4 0.038 0.011 0.982 1.138

5 29 27.8 1.18 285 34.9 1.123 0.807 0.71 12.2 0.067 0.11 0.511 1.0420 27.5 1.23 285 47.8 1.228 1.06 1 .198 15.2 0. 0 0.131 0.4 0.97221 28.3 1.11 284 88 . 1.57 1.726 1.032 13.9 0.063 0.011 1.20 1.07722 28.9 1.2 284 0.3 1.660 1.79 1.70 13.0 0.061 0.0171 1.270 1.03423 34.0 1.0 285 47.6 1.46 1.505 0.19 34.9 0.01 0.0121 0.8 1.0474 33.9 1.20 285 84.0 1.50219 .680 0.600 8.3 0.096 0.01 1.752 1.0168

3t 4 24.8 1.19 286 12.6 0.826 0.342 0.215 3.2' 0.0281,0)90( 0.150 1.205

25 2.4 1.08 285 1.3 0.854 0.47 0.919 2.5 0.060 1.07 0.115 1.2226 33.2 1.10 285 71.6 1.30 0 0. 16.585 2.3 0.012 0.015 1.26 1.0562 28.6 1.11 285 83.6 1.529 1.720 2.06 1.2 0.133 0.013 1.2 1.09435 28.5 1.08 284 8.3 1.58 1..7 1.70 20.3 0.103 0.020 1.235 1.067

29 28.4 1.051 285 4.1 1.240 .057 1.61 14.9 0.1098 0.161.0.481093130 27.6 1.0 1 2 86 3. 1.06 0.8I6 018420 1.6 0.016 0.1367 0.552 1.0684

31 24.8 1.10 288 12. 0.8262 0.342 0.215 31.2 0.137 1.010 0.04 1.205

32 24.5 1.10 286 15.2 0.856 0.40 0.1765 2.7 0.022 0.9928 0.29 1.1469

IA 33 34. 1.10 293 71.6 1.35 1.520 0.28 4.0 0.044 0.011 1.926 t 1.05 -

4 32.6 1.09 289 69.5 1.452 1.569 0.00 3.5 0.036 0.0102 1.672 1.026

35 28.5 1.10 285 87.1 1.577 1.744 0.746 20 0.038 0.0214 1.239 1.071_ -

36 4 28.3 1.12 2885 .94. 0.88301.857 1.26 19.6 0.038 0.9096 12 1.10 -

37 28.0 1.12 286 33.9 06 0.849 0.332 8.45 2.9 0.031 1017 0.540 1.054238 26.9 1.10 285 51.2 1.282 0.7 1.337 0. 167.2 46 0.1 3 0. 44 0.4 9 7 1.539 24.4 1.11 288 42.3 0.822 0.342 0.882 19.9 0.143 0.0139 0.41 1-92340 24.2 1.19 287 85.810.8593 10.415 016 2.33 . 02 0. 92 0. 2096 1.14 L.6

54 27.2 1.17 288 43.6 1.28 1.012 .65 .0.198 3. 0.011 0.410 10442

55 34.8 1.06 289 69.7 L.350 1.405 0.408 5.9 0.035 0.0127 1.76 1.071486 33.9 1 1.10 1 289 806.0 1.511 1.669 0.520 7.0 0.038 0.0111 1.660 1.0267

~135

Table B-2. Supplementary High Pressure Test Data (Concluded).

I [o -0g:0 -C

I I8II

i.29 ' 6 .5I 1.2 .5I1.4 1 . 9L. 0 28 2o

ac U- U~ a,

1.. 165 o16 2 0. 154 1.24. 69

58 23.0 1.122 M9 1 5 2 0.878 0 402 0.91 12.3 0.17 0.917 0.206 1.18 i 0.69359 27.0 1.178 294 34.3 1.136 0.800 1.33 0.150 0.518 1.053 0.843

60 26.5 1.140 291 47.7 1.28t 1.045 1.17 t0.31 0.450 0.98 o.835

16. 28.7 1.203 293 87.7 t.663 1.687 2.71 276 0.6 0.0198[ t.234 1.6 o9462 27.

" 1.162 289 103.3 1.780 1.917 4.14 3 0.01801 1.163 1

63 32 2 !.160 288 67.7 1.430 1.401 0.43 62 0 033 0.0122 1.5 108 114

64 1. 1.67 t 293 76.3 1.513 1.539 0.68 9. .5 0.0119 .77

1 A 65 1. 3. 0.l 7 1 185 93 f1

2 5 32.2 .34 292 .2 442 1.424 0.200

66 32.7 1.157 289 75.2 1.511 1.554 0.40 57 030 0.0128L.2 105 19

L72. .121 292 86.5 1.056 1.711 1.07 140 0.6 0.02021.3 107

68 2 7.9 1.123 293 102.5 1.793 1.949 1.28 162 0.7 0.0180 113 L00 .9

69 26.8 1.171 290 33.8 1.153 0.799 0).16 2. .1 0.14,1 0.3 .7 .5

70 26.5 1.076 20 47.8 1.291 1.069 0.31 4. 002 014 048 .o o85

71 24.4 1.085 24 14.6 1.082 0.357 0.13 17 017 0.920 0.9 116 076

1 72 1 23.4 1.057 29 2.0 0.912 0.321 0.16 2. 0.020 _ 0.948 . 16.5 1 '76

9A 73 22.9 1.030 23 14.8 0.900 0.389 0.31 4. 0.040 0.961 0.0 .0 .9

74 23.3 1.1o05 28 12.3 0.853 0.331 0.89 1 0.114 1.106 0.4 123068

75 27.4 1.061 27 33.2 1.100 0.790 0.85 1..5 0.076 0.147 0.2 107 o85

76 27.2 1.077 27 47.1 1.253 1.049 0.83 11.0 0.073 0.1210.7 095081

77 32.9 1.202 28 84.5 1.604 1.664 1.52 18.6 0.092 0.02541.5 103 093

78 28.8 1.199 23 102.2 1.771 1.920 2.98 29.3 0.18 0. 1 .081022.9

79 33.6 1.176 29 67.4 1.429 1.406 0.41 5. 0.031 0.0126 1.0 1.075 1.152

80 1 32.6 1.192 1 24 75.9 1.521 1.539 0.68 9 . 0 51 0 0110 1.8 1.057 1.5

8A 81 -33. 7 1.227 29 67.5 1.422 1.385 0.48 6. 0.T036 0.0123 .72 .07--" 1.158

C 8Z

82 33.7 1.260 22 76.9 1.520 L.529 0.71 96 0.054 0.OI12 1.36 1.043 1.163

83 28.3 1.197 25 88.6 1.638 1.699 M4.54 360 0.276 0.0192 1.243 1.074 0.994

84 1 28.1 1.212 29 Ot3.2 1.766 1:904 4.97 3. 0.299 0.0173 1.20 103 0.997

85 27.3 1.191 23 47.9 1.264 1.04 1.48 18 2 0.131 [0.126 04 6 0.981 0.846

86 26.2 1.134 22 34.3 1. L6 0.795 1.86 216 0.165 0.144 0 53 1.077 0.643

87 26 3 1.155 2 2 15.0 0:894 0.,88 0.47 6 . 06 0882 0 2 5 1,158 0.707

26. 0:061 0

88 25.1 1.133 12.4 0.85 . 0. 708

4A 89 22.8 1.159 28 12.6 0.856 0.328 0.93 1. 0.120 0.977 0.14 1.232 0.696

90 2.0 13.2 29 15.5 0.896 0.395 091 12.0 0.116 0.884 0.26 1.170 0.693

91 27.1 1.160 294 34.7 1.132 0.788 2.10 24.2 0.1186 0.153 0.51 1.045 0.840

92 26.7 1.178 291 48.8 1.269 1.043 2.71 27.5 0.240 0.124 0.4 0.979 i 0.84193 28.2 1.261 29 87.1 1.610 1.653 2.10 23.5 0.128 0.0189 1.24 1 E010.998

94 28.0 1.316 28 102.0 1.722 1.871 2.1 23.8 0.132 0.018 1.173 1.029 0.988

95 32.6 1.291 289 67.5 1.392 1.378 1.00 13.4 0.075 0.019 1.88 1. 077 1.1526 32.6 1.288 2 77.2 1.92 L.521 1..53 14.5 0.084 0.011 1.870 1.05 1.1559. 1.274 66.7 1.385 1.384 0.6 6.5 0.035 0. 1.50 -1.076 1.162

98 33.4 1.287 28 76.1 1.471 1.55 0.40 51.0 0.062 0.0110 1.874 1.041 1.163

99 27.8 1.261 29 86.2 1.585 1.672 3.24 10,5 0.197 0.001 1.3 1.09 0.986

S 28.0 1.224 23 102.7 1.734 1.905 2.29 248 0.1 0.0180 1.22 1.036 0.0

6O 27.1 1.15 20 34.3 1 127 0.788 L.48 18.3 0.141 0.18 0.504 1.061 0.837

102 26.2 1.161 288 48.1 1.261 1.05 01.75 20.7 1 0.155 0.116 0.481 0.90 0,850103 25.0 1.85 294 15.1 0.900 0.390 0.43 6.2 0.055 0.98 0.185 1.149 0.701

72 24.3 1.07 289 12.3 01.12. 0.32 t 2. 20.00 0.9 0.16 1.258.0.06

9A 73 22.9 1.248 289 66.1 1.382 1 3 94

0.26 4.8 0.020 0.010612.1 1.074 1.187

106 33.8 1.262 28 14.8 1.467 1.530 0.31 4.5 0.023 .10 2.030 1 1.042 1.190

107 28.9 1.247 288 85.3 1.55 1 679 1.50 18.5 0.091 0. 187 1 294 061 1.007

108 28.7 1.219 285 99.5 1.711 1.881 1.47 18.0 0.089 0.016 1.254 I 1.03 09

709 27.6 1.159 292 32.5 1.117 0.762 0.84 It.5 0. 75 1 0.18 21 1.063 0.851

790 33.6 1.1763 289 67.4 1.4290 1.40 0.41 1 5. 01 0.12 1.0 ) 07 5 1 .15

111 24.8 1.1952 294 15.1 0.903 0.396 0.17 2.5 0.022 1.026 0.17 1.141 0.693

112 26.0 1.135 794 12.3 0.858 0..931 0.20 3.0 0.026 1.054 0.142 1.210 0.63

136

87 .. ... . .........15 22 50 0 .89 .'8. 47.6..6 .82 025 1.5 .0

Table B-3. CO Emission Test Data Correlation.

(a) Data Correction to Engine Conditions:

Elco engine (EICO, test s 0.ni:

where:

s Vr) [0254)1.25 (9.42)0.5 24. [.exp 3

EICO, test g/kg EICO, engine, g/kgas measured, Table B-I corrected to SCO=

at test points indicated 1.000, 0.153, 0.0160 and 0.0072

Fuel Idle Cruise Takeoff Dash Idle Cruise Takeoff DashNumber (2) (3) (4) (5) (6) (7) (8) (9) (1.000) (0.153) (0.0160) (0.0072)

IA 68.4 60.9 8.5 7.3 1.3 1.3 1.0 0.9 63.3 8.5 1.1 0.61AR 75.5 75.2 7.4 7.0 1.2 1.2 0.9 0.9 71.0 7.8 1.0 0.52A1 91.8 80.9 9.5 8.1 1.3 1.3 0.8 1.0 85.2 9.0 1.1 0.63A 106.6 102.2 11.0 7.2 1.1 1.0 0.6 0.6 10C.4 10.1 0.9 0.44A 102.3 102.7 12.2 7.8 1.7 1.4 1.0 1.0 110.5 10.9 1.3 0.65A 92.5 98.5 8.9 8.1 1.3 1.5 0.9 0.9 92.1 10.1 1.2 0.66A 100.1 89.5 10.2 7.9 1.5 1.4 1.0 0.9 90.2 9.7 1.3 0.67A 113.1 101.4 13.9 7.3 1.3 1.2 1.0 1.0 104.2 11.6 1.2 0.68A 78.8 73.8 12.3 8.5 1.6 1.8 1.0 0.9 80.4 11.7 1.5 0.69A 94.5 86.2 11.1 8.7 2.5 1.7 1.1 0.9 87.6 11.3 1.6 0.610A 98.3 88.2 9.8 8.4 1.6 1.5 1.1 1.0 96.1 9.2 1.3 0.6IA 110.7 76.8 9.4 8.4 1.5 1.6 0.9 1.0 96.1 9.2 1.2 0.612A 93.7 82.6 9.6 6.9 1.5 1.3 0.9 1.0 94.3 9.9 1.2 0.613A1 103.0 99.4 14.0 10.9 1.8 1.9 1.2 1.1 100.9 12.9 1.5 0.7

(b) Data Correlation with Fuel Hydrogen Contentand Spray Droplet Size.

Engine Power Level Idle Cruise

b, Intercept 74.79 1 8.45m, Hydrogen Slope -0.977 -0.922n, Droplet Size Slope +0.482 +0.446r, Correlation Coefficient 0.740 j 0.756

137

Table B-4. HC Emission Test Data Correlation.


/Sco,engine21

EIHC, engine - (EIHC test) (SCO, 2.14

where:

25V -5 [ x 4 2" -T3

2O4_2 05412 (9.42)0. [e " 4 -T3)]

EIHC, test g/kg EIHC, engine, g/kgas measured, Table B-I / corrected to SCO

at test points indicated 1.000, 0.153, 0.0160 and 0.0072)

Fuel Idle Cruise Takeoff Dash Idle Cruise Takeoff Dash

Number (2) (3) (4) (5) (6) (7)- (8) (9) (1.000) (0.153) (0.0160) (0.0072)

IA 35.9 19.2 0 0 0 0 0 0 26.6 0 0 0IAR 34.5 27.8 0 0 0 0 0 0 27.2 0 0 02A1 41.4 31.9 0.6 0.3 0.2 0.1 0 0 35.5 0.5 0.1 03A 59.7 46.5 0.5 0.2 0.2 0.1 0.5 0.3 48.7 0.4 0.1 0.24A 49.2 38.1 0.8 0.4 0.3 0.1 0.2 0.1 50.7 0.7 0.2 0.15A 42.0 38.6 0.1 0.1 0 0 0 0 37.1 0.2 0 06A 45.5 32.1 0.8 0.4 0.2 0.3 0.4 0.2 34.5 0.6 0.2 0.17A 59.9 42.0 0.7 0.2 0.1 0.1 0.1 0.1 47.5 0.6 0.1 0SA 35.1 28.4 0.5 0.2 0,2 0.1 0.2 0.1 35.4 (1.

4 0.1 0.1

9A 45.8 44.1 1.6 0.4 0.4 0.1 0.2 0.1 42.5 1.2 0.1 0.110A 51.0 43.0 1.4 0.6 0.6 0.5 0.6 0,4 50.2 1.2 0.4 0.211A 50.6 24.2 0.2 0.2 0.1 0.1 0 0.1 38.8 0.2 0.1 0

12A 47.2 34.2 0.5 0.2 0.3 0.3 0.4 0.4 46.9 0.5 0.2 0.113A1 38.2 49.5 1.1 0.7 0.4 0.2 0.1 0.1 44.9 1.0 0.2 0

(b) Data Correlation with Fuel Hydrogen Content

and Spray Droplet Size.

IHC (T ( 5)m (4 -)"

Engine Power Level Idle Cruise

b, Intercept 32.05 0.0426m, Hydrcgen Slope -1.19 -13.2n, Droplet Size Slope +0.517 +3.44

r, Correlation Coefficient 0.597 0.744

138

Table B-5. NOX Emission Test Data Correlation.

(a) Data correction to Engine Conditions:

(ElSMOt, engine

11140,f engine - 1OX teat~s 0 ,ts

where:

(2. p3 \03 ~ -~ 6.29 h3

St4 *(2.6 k~w [*f) )ex, 192j 53.2 /and:

#I[f) 0.1243 f - 0.233, when f < 11.5

" 1.461 - 0.0231 f, when f > 11.5

ElNOx, test, g/kg EiNOx., engine, g/kg

as measured, Table B-1 corrected to SNOx

at test points indicated 0.168, 0.472, 1.000 and 1.816

Fuel Idle Cruise Takeoff Dash Idle Cruise Takeoff DashNumber (2) (3) (4) (5) (6) (7) (8) (9) (0.168) (0.472) (1.000) (1.816)

1A 2.1 1.2 6.1 5.8 14.6 13.9 23.2 22.0 1.8 5.8 11.6 23.31AR 1.7 1.4 6.6 6.4 14.4 13.7 22.0 21.5 1.9 6.6 12.0 23.72A1 3.7 3.6 6.6 5.9 13.8 13.5 22.6 22.6 3.7 6.3 11.7 23.63A 1.6 1.9 6.5 6.4 14.7 15.4 26.4 26.0 1.9 6.2 11.9 23.14A 1.6 1.5 6.3 6.2 16.2 15.4 24.5 23.8 1.3 6.2 13.1 23.55A 2.4 2.2 6.7 7.2 15.2 15.2 23.6 --- 2.3 6.4 12.3 23.36A 2.6 2.5 --- --- 16.0 13.8 -- 21.5 2.6 -- 11.6 21.97A 1.6 2.0 6.6 6.7 15.8 15.0 25.0 23.9 1.8 6.4 12.0 23.38A 2.1 2.4 6.7 6.3 15.2 14.9 25.6 25.1 2.2 6.2 12.3 25.39A 1.9 2.0 6.2 5.8 13.8 14.0 24.7 24.6 2.0 5.6 12.9 27.2

IOA 1.9 1.7 6.5 6.2 16.2 15.0 26.8 25.3 1.8 6.2 13.0 26.111A 3.2 3.1 6.5 5.8 --- -- --- --- 2.9 6.2 --. .12A 2.1 2.4 6.5 6.6 14.2 13.8 24.5 27.3 2.1 6.0 11.8 23.713A1 1.5 1.4 -- 6.2 15.9 14.0 21.8 25.4 1.5 6.6 13.2 26.3

(b) Data Correlation with Fuel Hydrogen Content:

EI H

Engine Power Level Idle Cruise Takeoff Dash

b, Intercept 2.220 6.18 11.80 23.65m, Slope +0.670 -0.034 -0.361 -0.195r, Correlation Coefficient +0.178 -0.051 -0.551 -0.228

139

Table B-6. Smoke Emission Test Data Correlation.


SN8 , test SN8, engine( From Table B-2 at (Average of two SN8 test\test points indicated) at each enzine condition)

Dash

Fuel Idle Cruise Takeoff Dash Takeoff 75% P3Number (2) (3) (4) (5) (6) (7) (8) (9) Idle Cruise (dry) dry

1A 3.0 0.7 1.8 0.8 8.8 10.5 3.3 3.0 1.9 2.3 9.7 3.2

IAR 4.0 1.8 2.6 3.0 17.0 17.1 5.9 8.0 2.9 2.8 17.1 7.02A1 4.1 2.0 8.4 8.7 10.3 16.0 3.1 4.3 3.1 8.6 13.2 3.73A 3.0 2.5 11.5 11.5 18.5 18.0 3.8 4.5 2.8 11.5 18.3 4.24A 12.4 12.0 24.2 27.5 23.5 23.8 13.4 14.5 12.2 25.9 23.7 14.05A 3.2 2.7 14.9 13.6 24.2 20.3 2.3 12.5 3.0 14.4 22.3 7.46A 17.0 4.4 10.2 15.2 13.9 13.5 3.0 8.3 10.7 12.7 13.7 5.77A 7.5 6.2 18.3 20.7 30.5 24.8 6.5 11.0 6.9 19.5 27.7 8.8SA 5.3 6.6 18.2 21.6 36.0 4.0 6.8 9.6 6.0 19.9 40.0 8.29A 12.0 4.7 11.5 11.0 18.6 29.3 5.8 9.3 8.4 11.3 24.0 7.6

IOA 16.5 12.3 16.8 15.0 27.6 34.4 6.2 9.5 14.4 15.9 31.0 7.91A 11.9 2.3 2.9 7.2 20.7 19.6 8.0 13.5 7.1 5.1 20.2 10.812A 1.7 2.2 2.3 4.7 14.0 16.2 3.0 5.7 2.0 3.5 15.1 4.413A1 3.8 3.8 16.6 19.8 28.5 24.8 4.8 6.2 3.8 18.2 26.7 5.5


SN b

Engine Poer Level Idle Cruise Takeoff Dash

b, Intercept 2.08 3.88 13.92 4.272, Slope -8.37 -8.93 -3.64 -4.06r, Correlation Coefficient -0.865 -0.800 -0.671 -0.680

140

4 O N a N - a WN CC 00 n a, 0', r- 0 C 0 a'D cm (0 m 0 m' IT 4 0 m WC C- a, Nc

m-m mt- 4N 0 m 0004 N 4 Ei4ONU

N~- 4V~~' v0~~' Wa N4 NN ' 0 400r N ' 4

- ~ ~ C 9CN 0 e0 N4 " 0N 4 N4'N 4C C4N 4 4' N0 4

N, Cm OtV 4 O 0 - 0N . N 0 0p Ch '' 1N a' 04 en0 N 0% '%eC a

C~4(e c~4% C N C-4 N

0 ~ ~ ~ ~ ~ C NN0 'Csr'- V4 NN .- M O0 '4 N 4 C,4 CN N O4

%- ANCCN 0' m' O4 '0 0 N, N44w n 4.NN NNN'%

N ~ ~ 4 NO, 0N O O 0' 0 N C 0ON a. N 10 rNO a

LC N4 C0~ 0 . 4 C 4 el N N C4 CI a' NO- f ~0

0 q. O" 0'^ W% m W% 0 w(% 0"0 0 0 GOm 0

U ' 44'0Oc" mm NN- ~~ N C N C. .4 -SOa'.'I? ON4 N

0h N0 NNNNN''""M4 -, 4A mm

N~e 06'~ ' C 4 N' CZ r4 e4 4 C Na

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mm mmo NNNN4444N-- n- NNNNN4 4 4 N- -4 4 N-

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w40 m K 1 Na'N' C4'CW 4'CON'C44 a'CO 0 - l% OIn a, aO 'C 4 W '~r-~1 NN MN N 10 ~ 4 44 ~ 4 4 N '

,a O N 4C 0~% 01N . 0%' 'a 0 OD 00OA..aNN CNC

N 4~4~"CCC- N4 04 C4 C4 r ~---NNNNN 4 4 C4 N CN 4 N4 N N4

C4 n V%0 rUW 0% 0 N enC00 fN ' 0 U- .O w 0 t, 'C, 00 0,SN~~~ 004a e44 U 04 4 N4 4 CC 'C 4n Nn colN' en es MUU cn el M4 I

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0 00 0 0 4 00UNIN0Im % 0 a100%0go tr4 c N Nri V 4 C4n CNN "nNNN N.' N2t"cn0 ntlr4t r4 g " I

a* O.~ O 0 N L N4 'n a, a 0 i 00C0M'D4%-- 4C 4N 0O U 44 (44 O 44. N C4 C- ' C N V~. N "4 4

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s0 N. N 04 N

-N M 44

00 00

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4 N M MM N 1 MC

.- ~~~ I . N~~I4VN N"" NN' II aa:!V 4

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N 0 M' NU M CF, 10 0 0 0 1 m M N cp1 Ch0 NC 0 0 M " '4 1'a M000

M) M MTMTr- 44 w m ~ 4 NC N0f' N 4@ O -7N

I0 04 on C4 M M on N C

E-v m CC N1 U 400 N 4 0' N N ' C0,0 M CN -U M M

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IN

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4 1

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N p-, P NOU. -~'-'

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M C4NN NN ~ N N N N

W4 N UNoo %AC4C N4

AN N NNo.s 4 .d,.4

t4 U .'~N0 NIUNNNN -~I

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IL 40 WN' U C 4 4 4 ~ ~ ~1_ NONtm O 4M en, N CqUP.

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M m 4 M 44 44 fnM M 4 r

143 0 0MN 4A.Q..N.C -.

Table B-8. Detailed Rear Liner, Transition Duct and Fuel Nozzle StemTemperature Data.

(T-T3), Temperature Rise, K, at Thermocouple NumberRear Liner Transition Fuel Nozzle Stem

Fuel ReadingNumber Number Avg. 26 27 28 29 30 31 32 33 34 35

2A 1 105 121 103 56 66 99 154 136 -44 -41 -392 86 85 91 50 61 92 122 105 -43 -39 -373 186 229 182 171 132 163 227 200 -74 -67 -654 242 329 234 189 186 207 301 254 -80 -72 -705 187 215 208 173 131 173 215 200 -69 -56 -586 219 268 237 218 154 189 241 227 -70 -58 -607 170 192 192 135 125 157 187 204 -101 -83 -878 189 221 207 162 139 174 200 221 -102 -84 -89 ,IA 9 162 193 182 114 Ill 157 177 204 -111 -93 -10010 177 212 191 139 124 165 194 216 -114 -95 -102

11 169 167 186 139 110 168 189 208 -79 -66 -7012 198 235 213 176 134 181 221 230 -78 -65 -6913 146 147 151 91 102 145 182 206 -82 -73 -7414 195 210 196 129 152 175 254 252 -87 -78 -80 ii15 98 95 103 47 68 100 132 142 -52 -47 -47S16 77 86 91 37 57 116 -52 -46 -45S-6A 17 -83 83 95 58 55' 77- 110 103 -44 -39 -3813is1 102 106 109 68 67 87 144 133 -48 -42 -4219 194 222 187 223 152 155 219 205 -84 -73 -7420 287 382 269 311 259 212 300 277 -86 -75 -77

21 196 232 224 199 135 164 212 208 -73 -60 -64

22 220 267 238 229 162 191 230 229 -76 -63 -6722 181 2L7 204 172 127 165 179 206 -109 -90 -9524 211 262 236 215 148 188 201 232 -110 -92 -9725 214 258 227 219 154 189 211 244 -104 -9 -9026 170 198 194 135 121 155 187 201 -102 -84 -8727 190 207 194 171 12 167 200 202 -71 -58 -6028 210 243 230 202 143 197 226 232 -72 -59 -6229 241 275 234 260 199 186 262 273 -89 -79 -8030 173 130 169 184 133 151 189 209 -87 -76 -7731 77 74 86 43 66 118 -47 -42 -414__96_9032 88 82 95 52 77 137 -49 -43 -431A 33 170 200 188 135 123 152 181 211 -114 -94 -9934 219 257 238 221 158 190 215 255 -118 -99 -10435 174 192 181 168 121 162 187 209 -75 -62 -6536 195 I225 203 197 139 178 207 222 -76 -63 -6637 148 147 147 119 116 135 171 206 -85 -75 -7538 235 255 222 2?3 204 192 272 283 -85 -77 -7939 75 70 81 37 62 76 85 119 -52 -47 -4640 96 90 102 37 83 90 j117 138 -54 -48 -48

144

Table B-8. Detailed Rear Liner, Transition Duct and Fuel Nozzle StemTemperature Data (Continued).

(T-T3), Temperature Rise, K, at Thermocouple NumberRear Liner Transition Fuel Nozzle Stem


II1AR 41 69 75 94 45 65 -48 -43 -42k

42 82 91 106 52 80 -50 -45 -41A43 160 169 175 112 133 131 223 181 -87 -77 -7744 201 225 209 134 179 161 283 220 -88 -79 -7945 205 241 241 182 150 178 239 210 -75 -64 -6546 181 201 214 161 125 158 217 194 -74 -62 -63

47 164 187 196 119 119 147 187 193 -112 -95 -9648 195 231 220 167 146 172 216 218 -111 -95 -96

13A 49 107 108 115 102 104 -38 -4050 83 81 96 68 87 -36 -3751 167 191 159 152 -63 -6452 213 239 215 187 -66 -6753 159 217 147 164 -53 -5354 170 239 187 179 -54 -5455 158 197 127 152 -75 -7756 164 198 133 161 -75 -77

10A 57 34 88 99 88 60 70 103 -3458 88 94 104 88 60 69 113 -4559 185 249 208 232 124 120 181 -6960 251 371 286 253 188 178 231 -7261 203 304 237 270 119 114 117 -5562 247 383 284 288 170 151 207 -5263 171 241 207 217 99 107 160 -81

64 199 265 225 270 127 130 182 -7912A 65 162 227 208 170 99 108 160 -74

66 177 243 214 204 109 115 180 -7767 179 243 211 223 115 116 169 -58b8 218 322 260 236 151 151 193 -5669 138 154 157 120 97 120 181 -6970 196 241 222 155 156 180 226 -7571 82 90 100 56 54 80 114 -4672 b7 75 84 45 42 65 95 -44

9A 73 94 100 109 86 55 81 134 -3374 8b 93 102 79 54 74 117 -2775 169 206 187 196 122 135 169 -62

76 236 309 261 227 190 205 227 -6877 151 196 185 175 96 128 130 -4678 241 332 264 274 189 194 194 -5579 159 0l 192 t82 96 134 149 -8780 201 265 223 260 122 160 179 -89

145

Table B-8. Detailed Rear Liner, Transition Duct and Fuel Nozzle StemTemperature Data (Concluded).

(T-T3), Temperature Rise, K, at Thermocouple Number

Rear Liner Transition Fuel Nozzle Stem


8A a1 164 231 202 163 103 139 149 -82 -8582 192 266 222 219 114 157 179 -84 -8983 209 256 214 304 139 167 177 -56 -5984 246 335 254 307 180 193 212 -57 -5885 264 321 261 317 232 213 241 -70 -7086 198 222 196 294 142 148 191 -65 -6387 101 Ill 116 118 62 82 117 -42 -4188 77 88 98 77 43 63 93 -43 -42

4A 89 100 119 134 94 57 87 Ill -3590 115 127 143 118 69 99 137 -7391 242 321 280 303 162 186 204 -6092 354 567 412 345 255 265 285 -70 -6893 222 287 252 262 144 188 201 -53 -5394 260 360 299 293 186 204 218 -56 -5595 165 250 231 179 113 159 178 -71 -74

E 1 96 205 270 251 224 133 167 189 -78 -847A 97 167 215 215 149 111 144 172 -78 -82

98 199 265 255 202 126 160 189 -79 -8099 216 292 262 245 131 164 202 -56 -58100 245 328 281 271 163 194 238 -55 -55101 197 224 219 242 133 153 214 -59 -63102 294 413 332 295 227 221 281 -71 -72103 94 103 126 72 52 76 140 -36 -37104 79 88 108 58 40 70 113 -36 -36105 167 225 231 139 106 136 168 -77 -81106 190 244 244 186 121 157 192 -82 -86107 190 221 223 222 122 160 1^3 -53 -56108 227 303 275 241 153 180 213 -56 -55109 166 190 195 172 109 129 201 -68 -t7110 233 292 273 216 192 188 241 -71 -70Ill 80 85 110 55 46 71 115 -39 -40112 71 79 100 51 40 b4 95 -35 -38

L I I

146

Table B-9. Liner Temperature Data Correlation at Takeoff (Point 7).

Data Correlation with Fuel

Hydrogen Content.AT - b + m (14.5-H)

Thermo-couple Installation I & 2 Installation 3Number a b r m b r Comments

1 6.81 144 0.706 16.98 249 0.236

2 16.36 276 0.527 Fuel IAR and 3A excluded

3 16.65 357 0.577- ---------- "stable film in Installation 3

4 21).15 400 0.841 41.09 321 0.976 Ut.atached only (IAIAR,SA, IIA,6A,3A,4A,7A,8A)

5 .. ..... ...- Unstable film

6 ... .---------- Unstable film

7 15.44 309 0.491 6.38 319 0.168

8 -3.70 404 0.102 22.73 410 0.633

9 ..........- 2.23 162 0.060 No Inst. l&2 data

10 5.6 197 0.293 33.32 196 0.570

II 7.58 471 0.238 7.58 471 0.238 Unattached both Installations (lA.2A,6A,9A,1OA,12A)

12 12.35 359 0.810 11.08 1N3 0.383

13 13.59 173 0.811 28.41 325 0.903

14 40.17 205 0.868 18.00 22110.863 Fuel ZA excluded

15 27.12 268 0.793 20.86 220 0.697

16 21.71 180 0.851 13.13 189 0.884

17 26.31 221 0.779 11.48 226 0.474

18 16.40 101 0.756 12.10 129 0.792

19 15 12 158 0.368 --- --- Insufficient Inst. 3 data

20 12.77 201 0.584 ------ No Inst. 3 data

21 17.31 182 0.927 17.26 201 0.478

22 16.56 96 0.872 0.72 137 0.028

23 6.00 93I --- ---- -- Fuels 1AR and IIA only24 9.44 109 0.438 14.33 131 0.931

26 5.43 241 0.309 22.56 300 0.769

27 1.54 227 0,097 5.92 264 0.339

28 15.84 184 0,770 28.84 224 0.978

29 8.40 143 0.439 12.86 149 0.749

30 3.37 181 0.413 9.09 166 0.376

31 3.59 127 0.284 5.09 202 0.297

32 3.50 221 0.429 - ------- --- Fuel 13A excluded, No Inst. 3 data

147

FAL_ _

A

Table B-10. Liner Temperature Correlation with Hydrogen Content(Thermocouple No. 7).

(a) Data Correction and Engine Conditions.

(TL-T3), Inner Liner Temperature Rise, K (TL-T3), Inner Liner Temperature(From Table B-7 at Test Points Indicated) Rise, K (Average of Two Readings)

Idle Cruise Takeoff Dash Idle Cruise Takeoff DashFuel

Number 2 3 4 5 6 7 8 9

2AI 78 .6 154 130 317 374 302 317 72 142 346 310

1A 90 88 155 148 278 294 295 303 89 152 286 299

6A 84 63 205 195 344 348 359 357 74 200 346 358

5A 78 98 276 239 336 328 335 325 88 258 332 330

V1A 91 105 191 203 317 327 339 358 98 197 322 349

lAR 88 70 123 92 299 290 296 303 79 108 295 300

13A1 69 75 227 192 324 321 340 354 72 210 323 347

10A 109 98 185 155 309 299 363 362 104 170 304 363

12A 84 95 172 155 303 283 339 307 90 164 293 323

9A 114 110 196 161 289 303 350 372 112 179 296 361

8A 121 146 289 269 333 346 371 354 134 279 340 363

4A 106 98 270 220 352 356 386 390 102 245 354 388

7A 98 106 283 197 348 351 350 377 102 240 350 364

3A 89 86 232 192 316 371 344 342 88 212 344 343

(b) Data Correlation With Fuel Hydrogen Content.

(TL-T3) - b + m (14.5-H)

Engine PowerLevel Idle Cruise Takeoff Dash

b, Intercept 80.5 147.8 304.9 306.6

m, Slope 9.0 35.2 13.0 25.9

r, Correlation 0.481 0.674 0.491 0.892

Coefficient

I

I

148

Table B-il. Flame Radiation Data Correlation.

(a) Correlation with Combustor Operating Conditions:

q R m SR + b

whereIP- 1.359\ 3

SR" exp 34.6 / + 73 -5

"" .9Qr, Radiant Heat Flux, kW/m2to / Calculated From Operating

o. 0.4V A Conditions Correlation at Sr$4 0 0 (0.696, 0.845, 1.000 and 1.180

g - Idle Cruise Takeoff DashP4 z~ * (0.696) (0.845) (1.000) (1.180)

IOA 8 113 5 0.984 83.9 100.7 118.2 138.6

12A 8 102 10 0.977 80.9 96.0 111.8 130.1

9A 7 83 23 0.931 80.8 93.1 105.9 120.8

8A 8 87 25 0.992 85.9 98.9 112.4 128.2

4A 8 79 29 0.994 84.3 96.0 108.3 122.5

7A 8 76 34 0.988 87.1 98.5 110.3 124.0

3A 8 62 44 0.953 86.8 96.0 105.6 116.8


= b + m (14.5 - H)

Engine Power Level Idle Cruise Takeoff Dash

b, Intercept 83.2 94.8 106.9 120.9

m, slope 0.60 1.32 2.05 2.93

r, correlation coefficient 0.208 0.470 0.421 0.366

149

Table B-12. Combustor Exit Temperature Profile Data.

Temperature Rise Ratio, ATLcI/ATAvg

Average Profile Peak ProfileFuel leading 1 2 3 4 5 1 2 3 4 5Number Number (Outer) (Inner) (Outer) (Inner)

2A 1 0.698 0.929 1.093 1.158 1.122 0.949 1.143 1.253 1.395 1.3312 0.682 0.929 1.066 1.178 1.145 0.942 1.168 1.189 1.415 1.3533 0.703 0.947 1.091 1.158 1.101 0.940 1.107 1.219 1.340 1.2414 0.735 1.003 1.083 1.109 1.070 0.989 1.283 1.303 1.330 1.2295 0.736 0.945 1.100 1.137 1.082 0.933 1. 174 1.304 1.356 1.2866 0.764 0.959 1.091 1.120 1.068 0.934 1.J99 1.314 1.349 1.2707 0.758 0.910 1.087 1.150 1.093 0.900 1.099 1.361 1.351 1.3238 0.778 0.922 1.034 1.156 1.109 0.907 1.109 --- 1.357 1.327

IA 9 0.757 0.913 1.063 1.160 1.107 0.881 1.099 1.365 1.33610 0.770 0.941 1.027 1.154 1.108 0.892 1.111 --- 1.362 1.32711 0.749 0.917 1.075 1.150 1.109 0.915 1.094 --- 1.341 1.30212 0.761 0.960 1.037 1.144 1.099 0.898 1.076 --- 1.349 1.29213 0.640 0.865 1.078 1.192 1.226 0.953 1.213 --- 1.352 1.41814 0.712 0.959 1.033 1.142 1.154 0.977 1.199 --- 1.339 1.30015 0.675 0.939 1.041 1.174 1.173 0.963 1.534 --- 1.343 1.35116 0.696 0.925 1.022 1.180 1.177 0.928 1.422 --- 1.388 1.365

6A 17 0.794 0.928 1.060 1.100 1.118 1.052 1.181 1.337 1.317 1.31218 0.803 0.930 1.063 1.097 1.107 1.111 1.190 1.318 1.287 1.32419 0.811 0.960 1.064 1.097 1.079 1.083 1.167 1.282 1.266 1.23820 0.830 0.981 1.072 1.082 1.034 1.098 1.180 1.262 1.226 1.17121 0.802 0.956 1.069 1.105 1.068 1.077 1.206 1.318 1.265 1.27522 0.809 0.966 1.066 1.099 1.060 1.101 1.210 1.311 1.296 1.24523 0.800 0.953 1.065 1.108 1.073 1.070 1.168 1.283 1.301 1.31124 0.811 0.964 1.044 1.106 1.075 1.100 1.192 1.300 1.316 1.305

5A 25 0.778 0.956 1.044 1.109 1.106 0.926 1.190 1.293 1.289 1.32926 0.770 0.937 1.070 1.121 1.109 1.074 1.191 1.307 1.275 1.34527 0.769 0.942 1.062 1.130 1.105 1.049 1.199 1.301 1.264 1.27928 0.807 0.967 0.980 1.135 1.111 1.056 1.203 1.301 1.291 1.29829 0.750 0.956 1.059 1.091 1.145 1.031 1.173 1.269 1.248 1.28430 0.764 0.934 1.062 1.075 1.164 1.012 1.153 1.277 1.257 1.32131 0.744 0.934 1.060 1.055 1.208 0.980 1.213 1.243 1.311 1.33432 0.725 0.907 1.049 1.106 1.212 0.921 1.191 1.216 1-.324 1.442

11A 33 0.792 0.972 1.033 1.103 1.093 1.086 1.186 1.294 1.285 1.33134 0.803 0.972 1.046 1.098 1.080 1.087 1.204 1.296 1.277 1.31235 0.781 0.941 1.057 1.117 1.105 1.006 1.147 1.267 1.244 1.26136 0.789 0.948 1.060 1.107 1.100 1.031 1.174 1.282 1.263 1.27137 0.742 0.922 1.053 1.109 1.173 0.943 1.147 .210 1.251 1.35538 0.778 0.938 1.049 1.107 1.127 1.000 1.124 ..218 1.241 1.24139 0.716 0.909 1.054 1.097 1.225 0.927 1.071 1.222 1.339 1.38040 0.750 0.931 1.068 1.079 1.172 0.959 1.203 1.230 1.310 t.341

lAR 41 0.796 0.961 1.025 1.101 1.118 1.029 1.168 1.303 1.293 1.29842 0.796 0.966 1.037 1.100 1.101 1.073 1.200 1.300 1.296 1.30743 0.872 0.972 1.053 1.095 1.059 1.117 1.222 1.319 1.278 1.29044 0.838 0.992 1.064 1.079 1.026 1.164 1.262 1.336 1.266 1.24345 0.833 0.964 1.076 1.093 1.043 1.136 1.267 1.353 1.297 1.26746 0.748 0.965 1.074 1.110 1.052 1.139 1.261 1.366 1.311 1.28547 0.785 0.960 1.085 1.111 1.059 1.119 1.237 1.353 1.334 1.33248 0.810 0.964 1.083 1.r99 1.043 1.133 1.253 1.351 1.317 1.312

13A1 49 0.791 0.934 1.066 1.115 1.095 0.996 1.219 1.226 1.291 1.26550 0.787 0.941 1.062 1.105 1.104 0.999 1.216 1.226 1.297 1.29051 0.869 0.924 1.084 1.090 1.033 1.125 1.241 1.242 1.153

52 0.785 0.969 1.137 1.086 1.022 1.107 1.212 1.300 1.237 1.15453 0.853 0.930 1.094 1.091 1.033 1.031 1.143 1.288 1.262 1.18854 0.859 0.931 1.104 1.086 1.020 1.059 1.133 1.265 1.237 1.14655 0.812 0.928 1.076 1.113 1.071 0.973 1.120 1.273 1.284 1.26156 0.827 0.927 1.073 1.109 1.064 0.986 1.079 1.280 1.294 1.266

150

Table B-12. Combustor Exit Temperature Profile Data (Concluded).

Temperature Rise Ratio, ATxocaI/ATAvg

Average Profile Peak ProfileFuel Reading 1 2 3 4 5 1 2 3 4 5

Number Number (Outer) (Inner) (Outer) (Inner)

IOA 57 0.810 0.990 1.042 1.133 1.026 1.040 1.351 1.447 1.283 1.14058 0.784 0.981 1.094 1.110 1.031 0.959 1.219 1.319 1.277 1.21959 0.866 0.995 1.101 1.070 0.968 0.935 1.180 1.243 1.196 1.11060 0.879 1.023 1.112 1.046 0.940 0.977 1.201 1.300 1.187 1.08961 0.874 1.002 1.095 1.067 0.963 1.020 1.127 1.206 1.263 1.17762 0.885 1.013 1.098 1.057 0.947 1.034 1.125 1.211 1.259 1.16063 0.913 0.974 1.076 1.064 0.972 0.983 1.108 1.159 1.223 1.16864 0.866 0.981 1.078 1.078 0.998 0.985 1.119 1.060 1.248 1.205

12A 65 0.860 0.986 1.085 1.076 0.993 0.978 1.118 1.161 1.231 1.186

66 0.867 0.993 1.085 1.071 0.985 0.981 1.120 1.185 1.246 1.19467 0.850 0.995 1.089 1.077 0.989 0.932 1.135 1.165 1.225 1.16668 0.868 1.013 1.090 1.065 0.963 0.964 1.157 1.181 1.227 1.13869 0.793 1.003 1.063 1.089 1.053 0.937 1.189 1.258 1.221 1.20270 0.830 0.993 1.092 1.073 1.013 0.958 1.173 1.228 1.179 1.11171 0.767 0.956 1.077 1.118 1.082 0.904 1.267 1.362 1.322 1.359

72 0.748 0.961 1.088 1.122 1.082 0.906 1.236 1.363 1.286 1.294I9A 73 0.791 0.969 1.077 1.097 1.066 0.866 1.181 1.311 1.223 1.201

74 0.819 0.983 1.067 1.088 1.044 0.940 1.276 1.256 1.202 1.1317 0.856 1.033 1.094 1.058 0.958 0.959 1.244 1.264 1.181 1.08976 0.910 1.015 1.109 1.041 0.926 1.120 1.276 1.276 1.188 1.035

77 0.857 1.019 1.094 1.068 0.963 0.977 1.099 1.208 1.210 1.13278 0.884 1.036 1.097 1.048 0.935 1.031 1.075 1.213 1.210 1.08279 0.861 1.012 1.078 1.068 0.982 0.957 1.074 1.181 1.213 1.130

80 0.874 1.018 1.076 1.058 0.975 0.981 1.088 1.182 1.715 1.151

M 81 0.871 1.013 1.072 1.063 0.981 1.016 1.170 1.262 1.263 1.17182 0.880 1.020 1.070 1.058 0.973 1.043 1.210 1.277 1.278 1.19083 0.859 1.019 1.082 1.062 0.977 0.975 1.091 1.185 1.197 1.118

84 0.858 1.017 1.083 1.063 0.979 0.986 1.092 1.186 1.208 1.12885 0.834 1.011 1.080 1.064 1.011 0.969 1.155 1.194 1.138 1.10886 0.805 0.994 1.080 1.085 1.036 0.927 1.162 1.221 1.155 1.14487 0.821 1.010 1.084 1.073 1.012 0.914 1.184 1.237 1.17G 1.11788 0.821 1.003 1.C93 1.071 1.012 0.887 1.185 1.246 1.183 1.088

4A 89 0.844 1,005 1.063 1.073 1.016 0.945 1.151 1.214 1.182 1.13690 0.836 0.999 1.069 1.074 1.024 0.909 1.173 1.234 1.171 1.13691 0.907 1.027 1.068 1.041 0.957 0.994 1.186 1.186 1.170 1.05992 0.911 1.031 1.068 1.034 0.956 1.054 1.173 1.187 1.142 1.07693 0.879 1.025 1.080 1.084 1.022 0.983 1.179 1.228 1.237 1.21794 0.862 1.016 1.065 1.064 0.994 0.958 1.129 1.180 1.197 1.16095 0.855 1.017 1.062 1.062 1.004 1.052 1.230 1.280 1.280 1.223

96 0.861 1.015 1.065 1.060 0.999 1.053 1.228 1.279 1.276 1.218

7A 97 0.844 1.006 1.063 1.062 1.002 1.029 1.206 1.277 1.288 1.23898 0.839 1.011 1.066 1.070 1.014 1.045 1.242 1.285 1.293 1.232

99 0.863 1.011 1.069 1.062 0.994 1.029 1.162 1.222 1.243 1.210100 0.862 1.008 1.065 1.063 1.005 1.020 1.171 1.221 1.240 1.208101 0.796 0.924 1.091 1.105 1.084 0.892 1.124 1.193 1.195 1.192102 0.863 0.970 1.077 1.064 1.026 0.940 1.092 1.146 1.176 1.205103 0.803 0.909 1.102 1 104 1.082 0.872 1.096 1.224 1.193 1.226104 0.785 0.932 1.124 1.086 1.073 0.910 1.141 1.489 1.183 1.274

3A 105 0.883 1.014 1.062 1.053 0.987 1.024 1.193 1.246 1.245 1.221

106 0.862 1.021 1.062 1.062 0.994 1.050 1.215 1.260 1.270 1.230107 0.853 0.999 1.068 1.072 1.010 0.958 1.165 1.214 1.223 1.194108 0.862 1.012 1.066 1.066 0.994 0.996 1.157 1.208 1.224 1.191109 0.779 0.980 1.089 1.092 1.059 0.843 1.020 1.179 1.166 1.178110 0.843 1.019 1.077 1.060 1.001 0,901 1.122 1.157 1.157 1.119

1ll 0.789 0.973 1.115 1.083 1.041 0.897 1.188 1.330 1.173 1.159112 0.801 0.978 1 136 1.073 1.012 0.881 1.187 1.536 1.151 1.133

151

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152

jl

APPENDIX C

CARBON DEPOSITION/EMISSION DATA

Two 24 hours tests were conducted to establish the carbun depositiontendencies of the J79-17C combustor. Fuels 1A (Repeat JP-4) and 13A1 (diesel)were used in these tests to represent the range of expected carbon depositionseverity. Both tests were begun with clean combustors and fuel nozzles.After each test, the combustor and fuel nozzles were visually inspected andthe fuel nozzle was flow calibrated. No changes in the fuel nozzle flowcharacteristics were detected after either test. The visual assessments ofthe combustor after each test are preseated in Table 14. Figures C-1 throughC-4 present the poattest photos of the tombustor liner and fuel nozzle.Limited repeat performance data was obtained during the endurance portion ofthese tests and a summary of pattern factor during this testing is presentedin Table C-1.

Table C-2 presents the detailed results fr6m the cascade impactor mea-surements on Fuels 4A and 10A.

1531

.2

2-..

r J.

F n I

I *2 IA

t

ct I

I 2

I

A

I

I

'9

qV

L. ___ U

Table C-I. Pattern Factor During 24 HourTests.

Test WC VT3Test Time, AT,

Fuel Point hr K P3 PF

1A 2 2.2 298 128 0.303(Repeat) 3 2.9 375 129 0.320

4 4.2 536 121 0.3195 5.7 678 124 0.336

7 7.8 591 117 0.3536 8.8 524 118 0.3668 10.8 452 134 0.3539 11.3 523 131 0.3518 12.0 434 132 0.376

8 13.0 438 132 0.3568 14.0 420 132 0.4097 15.0 539 118 0.3717 16.0 542 117 0.3876 17.0 512 117 0.3726 18.0 508 118 0.3785 19.0 644 123 0.2925 20.0 641 122 0.3174 21.5 487 126 0.288

3 23.0 344 126 0.355

13A 3 3.4 358 122 0.2912 4.0 276 126 0.2974 5.2 513 124 0.242

5 6.8 636 125 0.3006 8.6 510 121 0.2887 9.5 566 120 0.265

8 11.6 436 130 0.2849 11.9 444 130 0.2949 12.9 452 130 0.2669 13.8 468 129 0.3177 15.2 566 120 0.2925 16.0 727 122 0.1735 17.6 686 123 0.2254 18.1 496 127 0.2124 19.3 501 127 0.2023 22.6 291 129 0.3492 22.8 222 132 0.251

158 L

44..

I'

4 -4 -4 v4

u o u A~ A C O O Ot-

'a4 -4o4 N m I &M 1 co 1 0 4 C en 4 LA10 -ga)0

UU

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54[

0 I.4159

APPENDIX D

LOW PRESSURE TEST DATA

Two types of tests were conducted in the low pressure combustor test rig:altitude relight tests and cold-day ground start tests. Apparatus and pro-cedures which were used are described in Section V-B.

Detailed results for the altitude relight tests are presented in TablesD-1 through D-7. The combustor operating conditions are listed from whichthe simulated flight conditions were determined, and in the remarks column,the type of data point is indicated (LIGHT maximum altitude relight capabil-ity at normal minimum fuel flow rate, PBO pressure blowout, LLO leanlightoff, LBO - lean blowout).

Detailed results of the cold-day ground start tests are listed in TablesD-8 through D-11. At each combustor operating condition shown, lean lightoffand lean blowout fuel/air ratios were determined which are listed. Alllightoff attempts were successful, so each test was terminated only after theplanned minimum temperature (239 K) was reached.

F

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I

Table D-8. Ground Start Test Results, Fuel Numbers 1A Through 3A. 2

Combustor Operating Conditions Lean Blowout Lean Lightoff

wc AP Wf WfFuel TF T3 P3 (engine) P (engine) f (engine) fNo. K K kPa kg/s % g/s g/kg g/s g/kg

1A 227.5 227.5 101.8 3.175 0.90 8.8 2.77 21.2 6.67272.0 272.0 101.8 3.175 0.90 9.5 2.99 22.4 7.04266.4 266.4 101.7 3.175 0.88 10.5 3.30 21.2 6.67260.8 260.8 101.7 3.175 0.88 10.5 3.30 19.5 6.13255.3 255.3 101.7 3.175 0.84 10.5 3.30 22.9 7.20249.7 249.7 101.6 3.175 0.80 11.1 3.49 21.7 6.82244.2 244.2 101.5 3.175 0.77 11.1 3.49 21.7 6.82238.6 238.6 101.5 3.175 0.72 11.6 3.65 22.4 7.04

1A(R) 277.5 277.5 102.1 3.175 0.66 8.6 2.70 20.9 6.57272.0 272.0 102.1 3.175 0.65 9.2 2.89 21.4 6.73266.4 266.4 102.1 3.175 0.63 9.8 3.08 20.9 6.57260.8 260.8 102.0 3.175 0.60 10.1 3.18 19.2 6.04255.3 255.3 102.0 3.175 0.57 10.2 3.21 22.4 7.04249.7 249.7 101.9 3.175 0.52 10.8 3.40 23.1 7.26244.2 244.2 101.9 3.175 0.50 11.5 3.62 24.1 7.58238.6 238.6 101.8 3.175 0.45 11.7 3.68 25.3 7.96

2A 277.5 277.5 103.2 3.175 0.63 17.1 5.38 24.2 7.61272.0 272,0 103.2 3.175 0.62 18.3 5.75 23.3 7.33266.4 266.4 103.2 3.175 0.62 19.4 6.10 27.7 8.71260.8 260.8 103.2 3.175 0.55 21.2 6.67 25.2 7.92255.3 255.3 103.1 3.175 0.57 24.3 7.64 29.4 9.25249.7 249.7 103.1 3.175 0.53 25.6 8.05 33.1 10.41244.2 244.2 103.1 3.175 0.60 28.5 8.96 35.5 11.16238.6 238.6 103.0 3.175 0.54 26.5 8.33 33.1 10.41

3A 277.5 278.2 101.8 3.175 0.68 15.6 4.91 24.3 7.65270.4 272.0 101.8 3.175 0.64 22.4 7.05 25.0 7.86265.9 267.6 101.8 3.175 0.62 22.1 6.93 25.2 7.92261.5 261.5 101.7 3.175 0.61 23.1 7.25 26.5 8.32257-0 256.5 101.7 3.175 0.60 27.3 8.60 32.8 10.30249.7 249.7 101.7 3.175 0.59 27.6 8.68 33.6 10.60245.4 245.4 101.7 3.175 0.55 32.8 13.30 45.4 14.30239.8 237.6 101.7 3.175 0.54 39.6 12.40 75.9 23.90 =

168

Table D-9. Ground Start Test Results, Fuel Numbers 4A Through 7A.


Wc AP Wf Wf

Fuel TF T3 P3 (engine) P (engine) f (engine) fNo. K K kPa kg/s % g/s g/kg g/s g/kg

4A 277.5 277.0 101.1 3.175 0.68 18.3 5.75 29.9 9.40

272.6 271.5 101.1 3.175 0.65 17.3 5.44 27.6 8.68266.4 266.4 101.1 3.175 0.62 19.0 5.97 29.2 9.18260.4 260.8 101.0 3.175 0.62 21.0 6.60 31.1 9.78254.8 255.3 101.0 3.175 0.60 23.3 7.33 30.5 9.59249.6 249.7 101.0 3.175 0.59 29.5 9.28 41.8 13.14244.8 244.2 101.0 3.175 0.57 28.7 9.03 33.5 10.53239.3 239.3 101.0 3.175 0.51 32.8 10.31 46.0 14.47

5A 277.5 277.5 101.8 3.175 0.63 15.8 4.97 23.3 7.33272.0 272.0 101.8 3.175 0.61 17.6 5.53 23.1 7.26266.4 266.4 101.8 3.175 0.60 18.3 5.75 23.3 7.33260.8 260.8 101.7 3.175 0.60 20.2 6.35 23.4 7.36255.3 255.3 101,7 3.175 0.58 20.8 6.54 24.3 7.64249.7 249.7 101.6 3.175 0.49 21.2 6.67 26.7 8.40244.2 244.2 101.6 3.175 0.59 22.4 7.04 28.4 8.93238.6 238.6 101.6 3.175 0.56 23.2 7.30 29.7 9.34

6A 277.5 277.5 101.4 3.175 0.62 13.7 4.31 20.9 6.51272.0 272.0 101.4 3.175 0.60 14.5 4.56 20.8 6.54266.4 266.4 101.4 3.175 0.59 14.9 4.69 22.3 7.01260.8 260.8 101.4 3.175 0.59 15.4 4.84 22.6 7.11255.3 255.3 101.4 3.175 0.58 17.4 5.47 22.8 7.17249.7 249.7 101.4 3.175 0.58 18.3 5.75 25.3 7.96244.2 244.2 101.3 3.175 0.62 20.4 6.42 26.3 8.27238.6 238.6 101.3 3.175 0.48 21.8 6.86 27.2 8.55

7A 277.5 279.3 101.0 3.175 0.63 20.2 6.35 26.6 8.36270.4 273.2 100.9 3.175 0.60 18.3 5.75 27.3 8.58266.4 266.4 100.9 3.175 0.58 20.9 6.57 26.7 8.40260.8 262.0 100.9 3.175 0.58 22.7 7.14 27.8 8.74254.8 254.3 100.9 3.175 0.57 25.2 7.92 29.0 9.12249.7 248.7 100.9 3.175 0.57 25.2 7.92 36.4 11.40244.8 243.7 100.8 3.175 0.54 27.7 8-.71 37.0 11.60239.8 237.6 100.8 3.175 0.52 28.0 8.81 37.8 11.90

169

II

Table D-10. Ground Start Test Results, Fuel Numbers 8A Through 11A.


Wc p Wf WfFuel TF T3 P3 (engine) P (engine) f (engine) fNo. K K kPa kg/s % g/s g/kg g/s g/kg

8A 278.8 277.5 101.1 3.175 0.66 9.8 3.08 22.7 7.14272.6 272.6 101.1 3.175 0.65 8.1 2.55 23.6 7.42266.7 265.9 101.1 3.175 0.63 10.7 3.36 22.7 7.14260.3 261.7 101.0 3.175 0.62 10.1 3.18 22.7 7.14254.8 255.3 101.0 3.175 0.61 13.4 4.21 21.4 6.73249.7 250.0 101.0 3.175 0.57 13.9 4.37 21.5 6.76243.7 243.2 101.0 3.175 0.55 16.4 5.16 21.7 6.82238.1 237.6 100.9 3.175 0.54 17.0 5.35 22.1 6.95

9A 276.8 277.3 101.0 3.175 0.66 9.8 j.08 21.7 6.82271.0 272.1 101.0 3.175 0.65 8.1 2.55 21.4 6.73266.8 267.1 101.0 3.175 0.63 10.7 3.36 22.2 6.98261.2 261.2 101.0 3.175 0.61 10.1 3.18 23.7 7.45254.8 254.8 101.0 3.175 0.60 13.4 4.21 29.0 9.12249.2 248.8 100.9 3.175 0.60 13.9 4.37 27.7 8.71245.4 243.7 100.9 3.175 0.57 16.4 5.16 28.7 9.03238.4 238.4 100.9 3.175 0.55 17.0 5.35 29.4 9.25

10A 277.5 279.3 99.9 3.175 0.69 8.9 2.81 21.9 6.89271.5 270.9 99.9 3.175 0.67 9.2 2.89 22.8 7.17266.4 268.2 99.9 3.175 0.65 10.2 3.21 21.3 6.70261.5 262.6 99.8 3.175 0.63 11.0 3.45 23.9 7.53255.3 253.7 99.8 3.175 0.62 11.8 3.72 22.8 7.17249.7 248.7 99.8 3.175 0.59 12.2 3.84 22.3 7.01243.7 243.2 99.7 3.175 0.57 12.5 3.92 21.4 6.74237.0 237.6 99.7 3.175 0.57 12.6 3.96 22.3 7.17

IIA 277.5 277.5 101.7 3.175 0.66 8.9 I 2.79 19.0 j 5.97272.0 272.0 101.7 3.175 0.64 9.1 2.86 20.3 6.382 ).4 266.4 101.6 3.175 0.63 9.8 I 3.08 22.4 7.04260.8 260.8 102.8 3.175 0.61 10.1 3.18 23.7 7.45255.3 255.3 102.8 3.175 0.59 10.1 3.18 22.3 7.01249.7 249.7 102,8 3.175 0.59 12.0 3.77 23.3 7.33

244.2 244.2 102.7 3.175 0.58 12.3 3.87 21.8 6.86238.6 238.6 102.7 3.175 0.54 13.1 4.12 20.8 6.54

1 70

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4)171

APPENDIX E

FUEL NOZZLE FOULING TEST DATA

Fuel nozzle fouling tests were conducted in a small flame tunnel rigusing apparatus and procedures described in Section V-C. Primary results wereperiodic bench flow calibrations of the fuel nczzles to detect orifice plug-ging and/or flow divider valve seizure. These periodic flow calibrationdata, ordered by fuel type, nozzle inlet fuel temperature and accrued timeare presented in Table E-1.

172

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174***

APPENDIX F

SMOKE DATA CALCULATION

In this program, combustor component rig tests were conducted in whichsmoke emission levels were measured at the combustor exit plane by themethod specified in Reference 6. The result is a Smoke Number (SN) whichexpresses the opacity of filter paper that has been stained by the exhaustgases. SN is therefore, not a true thermodynamic property of the exhaustgas. A relationship between SN and carbon weight fraction (Xc), which is athermodynamic property, is presented in Reference 16. This relationship isreproduced in Figure F-I.

When combustor exhaust gaaes are diluted by turbine cooling air as theyare in the J79 engine, both SN and Xc are reduced. Smoke emission index (Els)g carbon/kg fuel, however, remains constant. Els is calculated by the rela-tionship:

ElS (xci) (io-0+

where:

i engine station where sample is taken

f = fuel/air weight ratio (g fuel/kg air)

Therefore, engine smoke level, which would be measured at engine Plane 8, canbe calculated from combustor rig measurements, taken at simulated engine Plane4, by the following procedure:

1. Measured (SN4) and (f4) at simulated engine test conditions

2. SN4 * (Xc4) (from Figure F-1)

3. ElS (Xc4) ((100

4. Cycle data * f8 at simulated engine test condition

5. Xc8 - EI s (0o0 f 8 (10-3)

6. Xc8 * SN8 (from Figure F-l)

For the J79-17C engine, f8/f4 - 0.838 at non-afterburning operating conditions.In the test data summary, SN4, Xc8, Els , and SN8 are all tabulted in TablesB-I and B-2 for possible future use.

175

~10 2

00

00

01 101

0

6g176

Io 0.......

0.0

E

101

0 10 20 30 40 50 60 70 80 90

Smoke Number (per SAE APII179)

Figure F-I. Experimental Relationship Between Smoke Number and ExhaustGas Carbon Concentration.

i 176

REFERENCES

1. Gleason, C.C., et al, "Evaluation of Fuel Character Effects on J79 EngineCombustion System," AFAPL-TR-79-2015, GE79AEG321, June 1979.

2. Gleason, C.C., et al, "Evaluation of Fuel Character Effects on FlO EngineCombustion System," AFAPL-TR-79-2018, GE79AE0405, June 1979.

3. Verdouw, A.J., "Fuel Character Effects on the Stoichiometric Combustor,"

AFAPL-TR-78-12, DDA EDR9356, October 1978, Confidential.

4. Jackson, T.A., "Fuel Character Effects on the J79 and F10 Engine Com-bustion Systems," NASA Symposium on Aircraft Research and Technology forFuture Fuels, Cleveland, Ohio, April 16-17, 1980, Paper No. 12, CP2146,

July 1980.

5. Longwell, J.P., Ed., "Jet Aircraft Hydrocarbon Fuels Technology," NASACP-2033, January 1978.

6. "Aircraft Gas Turbine Engine Exhaust Smoke Measurement," AerospaceRecommended Practice 1179, SAE, 1970.

7. "Procedure for the Continuous Sampling and Measurement of Gaseous Emis-sions from Aircraft Turbine Engines," Aerospace Recommended Practice1256, SAE, 1971.

8. Covey, M.W., "Final Report on Phase II of the Model J79-GE-1O SmokeReduction/Improved Life Engine Flight Evaluation," NATC SA-44R-79,August 1979.

9. Gleason, C.C., and Bahr, D.W., "Experimental Clean Combustor ProgramAlternate Fuels Addendum, Phase II Final Report," NASA CR-134972,January 1976.

10. Moses, C.A., and Naegeli, D.W., "Effects of High Availability Fuels onCombustor Properties," AFAPL-IOl, January 1978.

11. Blazowski, W.S., and Jackson, T.A., "Evaluation of Future Jet Fuel Com-bustion Characteristics," AFAPL-TR-77-93, July 1978.

12. "Aircraft Engine Emissions Catalog," Naval Environmental ProtectionSupport Service Report No. AESOl01-Revision 4, July 1975.

13. Topicz, A., "Low Smoke Long Life Combustion System," General ElectricEngineering Change Proposal No. 79U190, August 1974.

14. Colley, W.D., et al, "Development of Emissions Measurement Techniquefor Afterburning Turbine Engines: Supplement 2 - Afterburner PlumeComputer Program User's Manual," AFAPL-TR-75-52, Supplement 2, October1975.

177

15. El-Shanawany, M.S., and Lefebvre, A.H., "Airblast Atomization: The

Effect of Linear Scale on Mean Droplet Size," ASME paper No. 80-GT-74,

March 1980.

16. Shaffernocker, W.M., and Stanforth, C.M., "Smoke Measurement Techniques,"

SAE Paper No. 680346, April 1968.

178

*U.S.Oovernment Printing Office: 1981 - 757-002/348

V1 i

ad a095 057 - dtic · this final report is submitted by the general electric company, aircraft...

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