Study of advanced fuel system concepts for commercial aircraft

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Produced by the NASA Center for Aerospace Information (CASI)

; CR174751 - February 1985 RBSAEB166

-NI\S/\

National ~and C!"ace- ~ministration

Study of Advanced'Fuel System Concepts

for

Commercial Aircraft and Engines

" by - i ,j ,

George A. Coffinberry

General Electric Company ; 0" Aircraft Engine Business Group j' Clnclnnati,Ohio 45215 (NASl-CB-1747Sl) STUrY OE HVANC:',D fUEL .- 11€5-1~S78 SYSTE~ CCNCBp~s FCE CCbeEEe!!l AIECEAFT (General Electric cc.) 2!:E f Be A12/IIF ACl CSCI Ole

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G3/05 14346 prepared for

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National Aeronautics and Space Administration

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Lewis Research Center 21000 Brookpark Road - r-, Cleveland, Ohio 44135

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i , I. R'cort No. Recipient', Catalog No.

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NASA CR-174751 ..

Report o.te Ti'le and Subtitle 5.

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Januarv 1985 Study of Advanced Fuel System C,mcepts for Commercial Aircraft 6. PerfClrminq OrganizatIOn Codp.

and Engines

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i, 8. Performing Organization RePOrt No.

7. Authorhl R6~AEBI66 G.A. Coffinberry 10. Work Unit No.

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9. Performino O'9If'iZltion Name tnd Adm,.

Ceneral Electric Compan) 11, Contract 0' G,lnt ~.

I Aircraft Engine Business

Group rleumann Way NAS3-23267 Cincinnati. Ohio 45215 13. TYDe (1: ',eoort ",d Period Covered

'2. _no ........ _1IId_

Final National tleronautics and Space Administration 14. SponsoriftI;J Aglftey Code Lewis Research Center - 21000 Brookpark Road Cleveland, Ohio 44135 SuD'""*'1_V Notft '5.

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An analytical study va8 performed in ol."'der to assess n&lative pet"fot"m91\ce and. economic factors involved with alternative advanced fuel systems for future commercial aircraft

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operatins witb broadened p~operty fuels.

The OC-IO-30 wide-body tri-jet aircraft and the CF6-80X ansine vere used 8S 8 b~8eline design for the study.

Three advanced systems vere considered and vere spec~.fically aimed at add?:'"essins freezing point, thermal stability and lubricity fuel propcrti~a.

Actual DC~lO-30 routes an~ flisht profiles vere simulatod by , computer modeling and !.:'esulted in prediction Qf aircraft and engine fU61 system temperatures ., durins a nominal flight and durin! statisticJ11 , .• e-day-pcr-year cold and hot flights.

Emer&ency conditions were also evaluated. Fuel consumption and weight and ~awer extraction results vere obtained. An economic analysis VlB performed for new sit'craft and"-gystems, bssed. on present value and return on investm", .. t economic factors.

Advanced systl ~ means for fuel tank heat iRS included fuel red.rel'lation loops uiJing engine lube heat and generator heat.

Enviro~nental ~ontrol system bleeu air heat was us~ for tank heati", in a water recirculation lao? Advanced system eo,uponents included centrifugal fuel pumps and low pressure fuel nozzles.

The results showed 'hat fundamentally all of the three advanced systems are feastble but vary in their degree of compatibility with brcadened- property fuel.

Use of engine compressor bleed air for tank heating alons vi~h centrifugal I fuel pumps and low pressu~e n?z~!es yielded tho best performance and economic results.

L Ultl.m!.tely however, systems assign dee is ions would depend on the de&rea of antieipated fuel property change and other objeetives involving the design of the aireraft ~ engineo.

17. Key Wonk (Suqgnted by Authcwfl: I t8. Dimibvtion Sfotement

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Aircraft Fuel Future Fuel Aircraft Fuel System Fuel Freezing

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Thermal Stability Fuel Lubricity 19. S!curitv Ollulf. lof thil repa'U 22', ;~ice 20. Sc!euritv Clauif. lof It-is P89!1 No. of Page 21.

Unclassified • For sale by the National Technical Information S!!fvice. Snonp'field. Virglnl<1 22161 NASA·C-168 (Rev. 1OO7S)

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, " , ' FOREWORD This final report, documents work performed by the Advanced Technology Programs Department of the General Electric Aircraft Engine Business Group for the NASA-Lewis Research Center under Contract No. NAS3-23267. The General Electri.: Company wishes to express its appreciation for the keen interest and , technical guidance provided by the NASA Technical Manager, Mr. Stephen M.

Cohen.

J The General Electric Company Technical Program Hanager for this program l was George A. Coffinberry. Computer Modeling was performed by T. A. Hobbs.

The Administrative Program Manager was B.uno A. Alexander.

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section 1.0 SUMMARY

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2.0 INTRODUCTION

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3.0

J BASELINE AIRCRAFT AIlID ENGINE

3.1 Aircraft 3.2 Engine 4.0 ADVAIlICED SYSTEMS i ..

4.1 Introduction 4.2 Advanced System A J 4.3 Advanced System B o~. 4 Advanced System C 5.0 FUEL PROPERTIES 5.1 Fuel Property Control 5.2 Operational Consideration of Freezing Point 6.0 METHOD OF AIlIALYSIS AIlID FLIGHT DEFINITIONS " 6.1 Thermal Kodel 6.2 Kission Definitions i of 6.2.1' Airport Temperatures .~' 6.2.2 Commercial Forecasting Route Structure Analysis (CFRSA) 6.2.3 Environmental Analysis 6.2.4 Aircraft and Engine Analysis 6.2.5 Emergency Condition with Low Fuel Reserves 7.0 ACCURACY OF RESULTS 7.1 Introduction 7.2 Flight Ambient Temperature Effect 7.3 Energy Check 8.0 RESULTS OF FLIGHT STUDIES 8.1 Introduction 8.2 Flight Definition Results 8.3 Fuel Tank Temperatl .• res 8.4 Engine Temperatures r~ L.

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TABLE OF CONTENTS Section Page 9.0 SYSTEM EFFECT ON FUEL CONSUMPTION 199 9.1 Introduction 199 9.2 Results 200 10.0 ECONOKIC ANALYSIS 217 10.1 Introduction 217 10.2 Results 11.0 CONCLUSIONS 241 · .

12.0 REFERENCES J , , i , e'

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r LIST OF FIGURES

Number

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1 DC-IO Aircraft 8

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2 DC-IO Tank Fuel System 9 3 Auxiliary and No. 2 Tank Transfer to Main Tanks at Beginning of Flight 10

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4 Outboard Tank Fuel Transfer to Main Tank at End of Flight 5 Ft.'el Usage , 6 Fl. i.ght Engineer Fuel Panel 14 7 Fuel Transfer Schedule (Model Representation) 15

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8 DC-IO Pneumatic S~stem 16 9 CF6-80 Engine 10 CF6-80 Epgine on DC-10 18 Fu~l System Schematic 20 12 Fuel System Schematic 21 13 lOG Oil Cooling System 22 14 Fuel Pump 15 CF6-50 Fuel Nozzles (Pressure Atomizing) 24 16 CF6-50 Fuel Nozzles with Valves on Casing 17 Relationship of Advanced Systems to Fuel Properties 28 18 Baseline DCIO-30 and CF6-80X 29 19 Advanced System "A" 20 Advanced System "B" 31 21 Advanced System "C" 32 22 Weight Tradeoffs (Dry Weight, pounds/Kilograms) 33 23 Low dP Air Atomizing Nozzles 35 24 36 Nozzle 4P Characteristics 25 38 Nozzle Divider Valve Location= 26 Advanced System C 39 27 Remote Nozzle Divider Valve 42 28 centrifugal Fuel Pump (Gearbox Driven) 43 29 Aviation Fuel - Past Situation 46

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, LIST OF FIGURES (Continued) Number 30 Aviation Fuel Future Situation Increased Aviation Supply at Present Spec Limits 31 Freezing point Operational Options Freezing Point CUrrent Operation Requirements 55 Operational Planning for Lowest Fuel. Cost 56 34 Energy Transfer - Baseli'le 58 35 Energy Transfer - System A J 36 Energy Transfer - System B 60

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37 Energy Transfer - System C 61 38 Basic Thermodynamic Considerations for Tanks 63 39 Basic Thermodynamic Considerations for Components E = Total Thermal Energy (BTU) 64 J 40 Tank Heat Transfer to Air 65 · ~ 41 Tank Temperature Determination 67 42 Analysis Flowchart 69 · j ,; 43 Airport T~~eratures 74 44 Airline Route Structure Analysis Program , .

l 45 Wind temp Input 77 46 Nominal and Cold Flights

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47 Hot Flight 83 48 Cold Flight Air Temperature Based on NASA CR-135l98 86 49 Baseline - Cold Flight No. 1 Main Tank Temperature Based on NASA CR-135198 87 50 Baseline - Cold Flight No. 1 OUtboard Tank Temperature Based on NASA CR--135198. 83 51 System Energy Check - 1st Law of Thecmodynamics 52 System Energy Check - 1st Law of Thermodynamics Baseline - Nominal Flight Total System Energy Check 92 54 Energy Transfer Check (BTU/Kin) 55 System A - Hot Flight Fuel Pump and Lube HX Energy Check 56 Cold Flight Altitude Versus Distance , 96 · .

: 57 Cold Flight Altitude Versus Time vi

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LIST OF FIGURES (Continued)

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Number

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Cold Flight Mach Number 98

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lL 59 Cold Flight Installed Thrust 99 60 Cold Flight Engine Fuel Flow 100 , ' 61 Nominal Flight Altitude Versus Distance 101 62 Nominal Flight Altitude Versus Time 63 103 Nominal Flight Mach Number 64 Nominal Flight Installed Thrust 65 Nominal Flight Engine Fuel Flow 66 Hot Flight Altitude Versus Distance 67 Hot Flight Altitude Versus Time 107 68 Hot Flight Mach Number 108

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69 Het Flight Installed Thrust

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70 Hot Flight Engine Fuel Flow 110 Cold Flight Air Temperatures 112 72 Nominal Flight Air Temperatures 113 , '" 73 Hot Flight Air Temperatures 114 ; 74 Cold Flight Tank Fuel Level 115 75 Cold Flight No. 1 Main Tank Heat ~ransfer Area 116 76 Nominal Flight Tank Fuel Levels 118 77 lIot Flight Tank Fuel Levels 119 78 Cold Flight Auxiliary Tank Temperature (For All Systems) 120 79 Baseline - Co11 Flight No. 2 Main Tank Temperature 1:1 Baseline - Cold Flight No. 1 Main Tank Temperature 122 61 Baseline - Cold Flight No. 1 Outboard Tank Temperature 82 System A - Cold Flight No. 2 Main Tank Temperature 124 83 System A,- Cold Flight No. 1 Outboard Tank Temperature 125 84 System B - Cold Flight No. 1 Main Tank Temperature 126 85 System B - Cold Flight No. 1 Outboard Tank Temperature 127 1- 86 System C - Cold Flight No. 1 Main Tank Temperature 87 System C - Cold Flig'i,t No.1 Outboard Tank Temperature 129 vii fT," p;

"""~,,-""-"--"""-"""" -_'_T' ;":."~~~"" ~ •.• , ~"""!Ia'1)-,Ji -. q • LIST OF FIGURES (Continued) Number Title 88 Cold Flight ECS Bleed Air Flow 130 89 System C - Cold Flight ECS Bleed Air Temperatures and Precooler Setpoint 131 90 System C - Cold Flight No. 1 Main Tank Temp9rature 133 91 System C - Cold Flight No. 1 OUtboard Tank Temperatue Baseline - Hot Flight No. 1 Main Tank Temperature 136 93 Baseline - Hot Flight No. 1 OUtboard Tank Temperature 137 J 94 System A - Hot Flight No.' 1 Main Tank Temperature 138 l 95 System A - Hot Flight No. 1 OUtboard ~ank Temperature 139 96 System B Hot Flight No. 1 Main Tank Temperature 140 97 System B - Hot Flight No. 1 OUtboard Tank Temperature 141 System C - Hot Flight No. 1 Main Tank Temperature 142 99 System C Hot Flight No. 1 OUtboard Tank Temperature 143 , , 100 Baseline - Cold Flight Engine Fuel Temperatures 146 101 Baseline - ~old Flight Engine Lube and Scavenge Oil Temperature 147 102 System A - Cold Flight Engine Fuel Temperatu~es 148 103 System A - Cold Flight Engine Lube and Scavenge Oil Temperature 149 104 System B - Cold Flight Engine ~~e1 Temperature 150 105 System B - Cold Flight Engine Lube and Scavenge Oil Temperature 151 106 System C - Cold Flight Engine Fuel Temperatures 152 107 System C - Cold Flight Engine Lube and Scavenge Oil Temperature 153 108 System C - Cold Flight EngiM. Fuel Temperature 154 Baselin~ - Nominal Flight Engine Fuel Temperatures 158 System A - Nominal Flight Engine Fuel Temperatures 159 1 111 System B - Nominal Flight ~ngine Fuel Temperature 112 System C - Nominal Flight Engine Fuel Temperatures 161 113 Baseline - Hot Flight Engine Fuel Temperature 114 System A - Hot Flight Engine Fuel Temperatures 115 System B - Hot Flight Engine Fuel Temperatures 116 System C - Hot Flight Engine Fuel Temperatures 117 Fuel Residence Zones - Baseline Thermal Hodel

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• LIST OF FIGURES (Continued)

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11B Zone 1 Residence Time a~d Temperature - Nominal Flight 172 " 119 Zc'ne 2 Residence TiIr.,11 and Temperature - Nominal Flight 173 119a Zo.ne 2 Residence Time - Nominal Flight 120. Zone 3 Resid~nce Time and Temperature - Nominal Flight 175 120a Zone 3 Residence Time - Nominal Flight 176

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177 J' 121 Zone 4 Residence Time and Temperatura - Nominal Flight I 121a Zone 4 Residence Time - Nominal Flight 178 122 Zone 1 Residence Time and Temperature - Hot Flight 123 Zone 2 Residence Time and Temperature - Hot Flight 180 r- l23a Zone 2 Residence Time - Hot Flight 181 I l.

124 Zone 3 Residence Time and Temperature - Hot Flight 124a Zone 3 Residence Time - Hot Flight 183 125 Zone 4 Residence Time and Temperature - Hot Flight 11!4 125a Zone 4 Residence Time - Hot Flight 185 126 Base1in~ Emergency Flight - Distance 187 127 Baseline Emergency Flight - Altitude 188 12B Baseline Emergency Flight Engine Fuel Flow 189 129 Baseline Emergency Flight - Main Fuel Tank Quantity 190 Baseline Emergency Flight - Outboard Tank Fuel Temp~rature J91 131 Baseline Emergency Flight - Engine Fuel Temperatures 192 132 Emergency Flight System A - OUtboard Tank Fuel Temperature 193 133 Emergency Flight System A Engine Fuel Temperatures 194 134 Emergency Flight System B - OUtboard Tank Fuel Temperature 135 Emergency Flight Syst~ B - Engine Fuel. Temperatures 196 136 Emergency Flight System C - OUtboard Tank Fuel Temperature 197 137 Emergency Flight System C - Engine Fuel Temperatures 138 Nominal Flight ~SFC for 115 lb. (512 N) Weight Baseline.

Systems A and B 201 139 System C - Nominal Flight ~SFC for lS5 lb. (S23 N) Weight 202 140 Baseline - Nominal Flight Fuel Pump Input Horsepower 203 ix

LIST OF FIGURES (Continued).

Number Title Page 141 Baseline - Nominal Flight ~SFe for Fuel Pump HP 204 142 System e - Nominal Flight Fuel Pump Input Horsepower 205 143 System e - Nominal Flight ~SFe for Fuel Pump HP 206 144 Nominal Flight EeS Bleed Air Flow 207 145 Nominal Flight ~SFe for Ees Precooler Fan Bleed Baseline, Systems A and B 208 Baseline - Nominal Flight 4Fue1 Burn for EeS Fan Bleed. Fuel Pump HP. 115 lb. (512 N) Weight 210 System A - Nominal Flight 4Fuel Burn for EeS Fan Bleed. Fuel Pump HP. 115 lb. (512 N) Weight 211 148 System B - Nominal Flight 4Fue1 Burn for EeS Fan Bleed. Fuel Pump HP. 115 lb (512 N) Weight 212 149 System e - Nominal Flight 4Fue1 Burn for EeS Fan Bleed. Fuel Pump HP. 185 lb. (823 N) Weight" 213 150 System A Economic Factors Per Engine Jet A 229 i 151 System B Economic Factors Per Engine - Jet A 230 152 System e Economic Factors Per Engine - Jet A 231 153 System A Economic Per Study Fuel Fac!:ors Engine - 232 154 System B Economic Factors Per Engi.ne Study Fuel 1-0-·4 J 155 System e Economic Factors Per Study Fuel Engine - 234 System B Economic Factors Per Engine - Study Fuel Fuel Cost for Zero Investment Incentive 235

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LIS'!: OF TABLES

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Table No,

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Estimat~d Fuel Processing Requirements 49

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2 Fuel Property Specification Limits 50 3 Approach to Fuel Property Issues S2 4 Tank Heat Transfer to Air 66 5 Defining Flights 70 6 Obtaining Cycle Data 72 ..

7 Selected Flights BO Flight Temperatures 81 Cold Flight Tank Temperatures Minimum During Flight _ OF (OC) 9 135 Nominal Flight Tank Temperatures Minimum During Flight - OF (OCl 144 11 Hot Flight Tank Temperatures Minimum During Flight - OF (OC) 145 12 Fuel and Engine Lube Oil Temperature Minimums During Flight OF (OC) \ ; Main Fuel Pump Inlet Temperature Maximum During Flight - OF COCl 166 14 Fuel Control (KECl Fuels Temperature Maximum During Flight - .~ OF (OCl 15 Nozzle Fuel Inlet Temperature Maximum During Flight _ OF (OCl 168 16 Engine Scavenge oil Temperature Maximum During Flight - ~¥ (OC) 169 17 Nominal Flight ~ Block Fuel for Pump Power, Fan Bleed, and Weight (Loss) Per En~ine 214 .1 1- 18 Nominal Flight ~ Block Fuel for Combustor Fuel Preheating (Gain) Per Engine 19 Nominal Flight ~ Block Fuel Net Including Loss and Fuel Heating ~ain Per Engine 216 i.

20 Economic Influences 219 I 21 System Cost to Airline Per Engine 221 22 Life Estimates 23 Economic Tradeoffs - Future Fuel 227 24 Economic Tradeoffs - Jet-A Fuel 228

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NOMENCLATURE AND ABBREVIATIONS AERO Aerodynamic AIC Aircraft APU Auxiliary Power Unit AUX Auxiliary AVG Average BLD Bleed BTU British Thermal Unit o C Degrees Celsius Commercial Aircraft Mission Analysis CAHAL Capacity CAP !

CDP Compressor Discharge Pressure , c CFRSA Commercial Forecasting Route Structure Analysis COMP Compartment .'

Cp Specific Heat cst Centisi:.okes DAC Douglas Aircraft Company i dc Direct Current DOC Direct Operating Costs Delta (Change in) Specific Fuel Consumption DSFC Energy E Environmental Control System ECS Engineering ENG'G F Filter ) f( Function of ( ) o F Degrees Fahrenheit Federal Aviation Regulation FAR FP Freezing Point ft Feet Feet ft Cubic Feet Gal Gallon GIB Gearbox

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'.~ " NOMENCLATURE AND ABBREVIATIONS (Continued) I

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,i GE General Electric Company GNP Gross National Product GPM Gallons Per Minute HP High Pressure hp Horsepower hr Hour HX Heat Exchanger IDG Integrated D~ive Generator IFR Federal Aviation Regulation International Fuel Reserves < ISA International Standard Atmosphere k Thousand

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K Degrees Kelvin

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kg Kilogram kJ Kilojoule kVA Kilovolt Amperes e, " lbs Pounds

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I M Mass \ M Hotor , Max Haximwn I j.

J " !lEC Hain Engine Control MFF Hain Fuel Filter 1:IFP Main Fuel Pump HHX Main Heat Ex~hanger Min Minimwn min Minutes Mp Mach Number NM Nautical. Miles IfMI Nautical Miles

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No, Number OB Outboard

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NOMENCLATURE AND AllBREVIATIONS (Conti/.ued)

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; . t· AP Differential Pressure pph Pounds Per Hour ppm Pounds Per Minute pps Pounds Per Second press Pressure psi Pounds Per Square Inch Q Heat Lube Heat Rejection Q L 'I , Pump Heat Rejection Q p QTY Quantity REG Regulator ROR Rate of Return RPM Revolutions Per Minute sfc Specific Fuel Consumption S/O Shutoff I , SPEC Specification i SST Supersonic Transport SYS System SYST System T Temperature AT Differential Temperature Air Temperature TA Air Film Temperature

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A Ambient Temperature TAHB TBC The Boeing Company TEMP Temperature Recovery Temperature TR TRANS Transfer Wall Temperature Tw Air Total Temperature T2 U Air Film Coefficient UCR Unscheduled Component Removal Rate xv

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~ NOMENCLAtuRE AND ABBREVIATIONS (Concluded) Velocity Vor. Vo1ume W I'lass Flow WFE Engine Metered Fuel Fl('if (Weight) WIreS Waste Heat Recovery System W,T Weight SYMBOL FREQUENTLY USED: Throttling Process xvi

{~' 1. 0 SUKMAR'!

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A computer model representation was made of the DC-IO-30 aircraft fuel .

system and the fuel. lubrication. bleed air. and generator cooling systems for

1- the CF6-80X engine. This model was derived from engine thermal models for

CF6-80A/Al engines used on th~ B-767 and A-310. The thermal model was used in conjunction with General Electric Commercial Zngine mission analysis computer codes to project and simulate actual routes and flight profiles for the DC-lO-30. The flights included nominal and statistical one-day-per-year cold and hot extremes.

With the DC-IO-30 and CF6-80X serving as a baseline. three new advanced fuel systems were evaluated under the same simulated flight conditions in order to assess differences in performance. The primary objective of this , , work was to determine feasibility and relative degree of compatibility of these alternative systems with future broadened-property fuels. The fuel ' properties determined to be of primary interest were freezing point. thermal stability and lubricity. An emergency flight condition representing low ruel , " reserves was also evaluated.

:.i I In addition to system performance considerations. fuel consumption and J economic tradeoffs were also addressed. Component weight, engine air bleed ,'i and shaft power extraction. and the effect of fuel heating were determined in order to arrive at relative differences in fuel consumption for the different sy~tems. These results along with estimates of acquistion cost and J maintenance cost were used in an economic analysis which yielded differences in economic incentive for e,ch sy~tem.

The three advanced systems included means for aircraft tank heating by energy extraction from existing engine systems. These heat sources included engine lubricating oil. generator oil. and engine compressor bleed air for the aircraft environmental control syste~. Fuel recirculation to the tanks was used for the engine lube and generator oil approach. A water recirculation loop was used with the bleed air system. The advanced systems also included

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• alternative fuel pump designs for increased tolerance to low fuel lubricity.

A centrifuga.1. pump was used with the most advanced system. Fuel nozzle configuration:; to increase tolerance to low therntal stability fuel properties included air atomization and alternative nozzle divide,r valve designs.

The reButs of the study showed that all systems are generally feasible but vary in dp-gree of compatibility with broadened-property fuels. The coldest bulk fuel temperature predicted for the main wing tanks was -45° C (-49· F) for the baseline design. This is 5· C (9° F) less than thp. maximum freezing point cor Jet-A fuel indicating the probable use of Jet-B or Jet-Al by the airline for this flight. These results during the same cold flight improved to -33.9° C (-29° F) using engine lube oil heat and -41.7° C (-43° F) J : using generator oil heat. Bleed air heat would have resulted in a -32.2° C (-26° F) tank temperature if the precooler setpoint was 121° C (250° Fl. Thus the bleed air system affords a 12.8° C (23° F) increase in fuel freezing point. The only significant problem revealed by the study was the use of engine lube oil heat for fuel freezing protection. Engine oil temperature might become too cold fo.r proper distribution in the. engine. Oversizing the oil-to-fuel heat exchanger (degrading performance) might be a solution but is not desirable from a weight standpoint.

Fuel lubricity problems caused by severe hydrotreating ~o~l~ be solved with the use of new pump gear materials or by changing to centrifugal pumps.

It is highly uncertain. however. that low lubricity will be a fuel problem in ~. - any event.

Thermal stability for the study fuel was estimated to be 220° C (428° F) minimum (breakpoint) as compared to 245° C (473° F) minimum for Jet-A. This reduction would increase projected unscheduled component removal (UCR) rate 6 6 from 50 per 10 hours to 140 per 10 hours using the study fuel. The worst of the advanced systems nozzle concepts yielded a UCR rate of 41 per 10 hours using the study (broadened-property) fuel. Hence from a fuel nozzle standpoint. it is believed that thermal stability reduction could be accommodated with a new design.

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·." •••• ~~- .•..••• - ••• "p~ "~~-~' -'"'"''''''~'-"'-'''='"'''''''t;)'l!II4r-:'1 • The emergency flight simulations with low fuel reserves and continuation of tank heating did not reveal major problems. With relatively simple provisions the fuel tanks would not be subject to overheating.

The economic results showed that only a small improvement is needed in fuel consumption to offset the initial cost of the most complex system. Using ~ fuel as a heat sink for engine bleed air eliminated engine fan bleed cooling L (precooling) and resulted in a 0.312 percent improvement in block fuel consumption for the nominal' flight. After consideration of initial investment l~ ana maintenance cost differentials. the net 1982-dollar investment incentive for this system was $93,669 ($69.254 with study fuels) per engine system over the 15-year life of the aircraft.

An additional result gained from this program was the apparent benefit of system computer modeling. Advanced systems and flight simulations are far too complex to afford assessment in a simple manner. The relative differences are small. In future work. this type of approach may be of absolute necessity if the best designs and operational decisions are to be made.

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2.0 INTRODUCTION This final reprr.t presents the results of a study performed by the General Electric Company Aircraft Engine Business Group, under NASA Contract NAS3-23267 for the NASA-Lewis Research Center. The study involved computer modeling of the fuel system for a wide-body commercial aircraft using high bypass jet engines. Advanced fuel systems were compared with a baseline design on the basis of similar flight conditions in order to ascertain compatibility with fuel properties and economic impact to the aircraft user.

In the future it can be expected that there will be a general reduction in the availability of petroleum crude sources. This, coupled with a relative increase in aviation transportation and increased competition for kerosene-type fuel may lead to an incentive for the use of broadened-property aviation jet fuel. In the aircraft and engine fuel system, fuel property change may influence tank fuel freezing protection, combustor fuel nozzle deposits related to thermal stability, and engine fuel pump wear resulting '.'

from low lubricity fuel characteristics.

It is recognized that the design evoluti"", of aircraft and engine fuel .~.

systems should anti!!ipate such changes to enable the aviation industry to provide a reliable·and cost-effective product. The objective or-this program was therefore to define and analyze a variety of advanced fuel systems which address these fuel property issues. Performance of the advanced systems relative to fuel compatibility, aircraft safety and economic benefits was the key objective.

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3.0 BASELINE AIRCRAFT AliD ENGINE

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The MacDonnell Douglas DC-10-30 wide-body tri-jet aircraft was chosen for the study because it represents a size and range which is of particular

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interest from the standpoint ~f low temperature fuel. Figure 1 shows the

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DC-10-30 aircraft. The DC-IO fuel system involves fuel transfer from several , , tanks to a particular engine during the course of a prolonged flight. Thus, the combination of prolonged fuel cooldown time and a modern tank-to-engine fuel management system provided an excellent baseline for this study of advanced systems involving tank heating.

Other particulars of the DC-10-30 are as follows: range 11,112 tan 6,000 nmi 256,284 kg 565,000 Ibs max gross take-off weight 45,360 kg 100,000 lbs max payload seat passenger load i The DC-10-30 fuel system includes eight fuel tank compartments. Triple or quadruple redundant tank boost pumps for all tanks and all-attitude suction-feed for wing engines is provided. Figure 2 shows the overall fuel , .'

management system.

For the study, only the No. 1 engine (left wing) was considered.

~eferring to the simplified fuel system in Figure 3, fuel is transferred by boost pumps from the auxiliary (center) tank and No. 2 (tail engine) tank to the main outboard tank, then to the main tank. This occurs at the beginning ..

of the flight in order to burn off fuel in the center of the aircraft. This L.

is desirable on large aircraft from the standpoint of wing structural load and for aerodynamic stabil\ty. As the flight progresses, fuel weight distribution shifts toward the outward fuel tanks. At the end of the flight, fuel is t~ansferred from the main outboard tanks to the main tanks as shown in

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ELECTRICAL CONTROL Figure 2. DC-lO Tank Fuel System.

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--- No. 2 Tank Fill Valve

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No.1 ----.

• .. ,1 Tank , Fill Valve 'j " I ) Aux.

Tanks No.1

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No.1 t

Main OB

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Tank iTanks Transfer

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Pump

I No. 21

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Tank

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L-Overflil .,i.

Drain • Fuel is Transferred When Transfer Pumps are Turned ~ , On and Fill Valve. are Opened Illy Flight ~ngineer ~ !l • Relief Valves (Not Shown) Prevent Overpressurizatlon

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j: Figure 3. Auxiliary and No. 2 Tank Transfer to Main

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;i Tanks at Beginning of Flight.

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Figure 4. This occurs automatically on the nC-IO if the main tank level is low enough that it becomes desirable to consolidate all remaining fuel reserves in the main tanks which are directly feeding the engines. Figure 5 I.

shows sequentially how fuel is transferred and burned during a long-range flight extended to minimum fuel rese~es. Figure 6 shows the fl~ght engineer's pane1 and the fuel transfer/burn selections Which can be made.

The computer model used for this study included the fuel transfer and fuel burn aspects discussed previously and shown in the figures. Figure 7 further describes the fuel transrer schedules used in the model.

Advanced System C uses engine compressor bleed air as a h~at source for tank heating. This bleed air is normally used on the DC-lO for wing anti- icing and for the cabin environmental control system (ECS). Figure 8 showa the baseline DC-IO plleumatic system associated with this bleed air. Note the location of the precooler setpoint which affects the amount of heat removed from the engine bleed air.

3.2 ENGINE The General Electric engine currently used on the DC-IO-30 is the CF6-50C. For the study, however, a newer engine, the CF6-BOX, was modeled.

The C~6-80X is a potential growth version of the CFG-80 used on the B-767 and A-310. The -80X engine would have about 6 percent better cruise sfc than the

-soc. For the installed engine, standard day sea level static rating would be

as follQ~: 56,000 249.1 kill lb. thrust 4.1:1 fan bypass i 131.5 kg/s 290 pps core airflow !

, .

8890.4 kg/h 19,600 pph fuel flow

I~

The CF6-80X engine is shown in Figure 9. The engine installation on the DC-IO is shown in Figure 10. The engine fuel system is shown schematically in

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'f Continuous From Boost , FueU Makeup '~, Pumps Float " for Water Valve , Scavenging

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' Trans SIO Outboard Valve

! Main Tank!

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Float Valve t: ' ....

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Jet Pump 'j, Je# Pump Trans ~ Trans ~j\ 1'/;.

t Baffle (Flow Inward Only) "~ j

• low Fuel Level In Main ~ank Causes Float Valves to Automatically

Open Outboard Compartment Grav!ty Drain Valve and Activates Jet Transfer Pumps - Transferring Fuel to Inward Area of MaIn Tank Regardless of Aircraft Maneuver Outboard Tank Fuel Trans fer Figure 4.

To Main Tank at End 0 f F1 ight.

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Aux Auxiliary 777% ~ Transfer If { ( ! ! < ( < a , ,evel Decreasing Full (35.292 Ib Cap.)

(98.220 Ib Cap.) No. 3 Engine 4 ~

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,~ Transfer Full Empty , Engine 2 ,.

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Transfer F.inpty Level Decreasing evel Decruaslng Empty j Engine 2 Figure 5. Fuel Usage.

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Fuel Used From En~ine

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/ ~.. . '- '-.: ' , '~ P:., ~ • ., / .. '"

___ Fuel Temp Sensor l

Located On Front

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Spare of No. a

Outboard Tank ~ X-Feed -!~~~'lFli:::: NormalUYf1 : t' Closed I 1"'-::;:';: 0 , .....

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~ Transfer Pump 111' o.? ~ 'i' [J

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iA On With , ...... , ....

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--- Open With ; j Fuel Open With Auxiliary and I; Open With Qly From , Auxiliary and ., No.2 Transfer Auxiliary ;J Tank Level No. 2 Transfer ~ ; Transfer t; Sensors !

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Figure 6. Flight Engineer Fuel Panel.

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CD Auxiliary Tank Transfer

• Used on Cold Flight Only (long Range) • Starts at Beginning of Flight • Rate Equals Engine Burn Rate • No Change in Other Tank leveDs (Full) , .1 • From Aux Equals 98,220 ibs (44,550 kg) (Full) to 400 Rbs (181 kg) (Residual)

,-

CD No.2 Tank Transfer

• Starts When No.1 or No.3 Mains Plus Outboards at 21,000 Ibs (12,250 kg) io 33,110 Rbs (15,320 kg) " ....

• Rade of 25,200 pph (3.2 kg/s) to No.1 and 25,200 pph (3.2 kg/s) to No.3 IJ> ~, • Main Tank leveDs Increase - No Chall'llge In Outboard Levels (Full) j, ~~ • Slops When 1N0. 2 Tank 500 Ibs (227 kg) More Than ADD Other Tanks Combined It ,

CD Outboard Transfer

.~ + • Starls When Main Tank Level Goes BeDow 5000 Dbs (2210 kg) • i • Rate Equals Engine Bum Rate -, • No Changes in Main ~~lfeg «5000 Ibs) «2210 kg) • Stops When Outboard Goes Below 400 Ibs (181 kg) (ReSidual) Figure 7. Fuel Transfer Schedule.

(Model Representation) (;v.

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FANENGIN~ ... __ CORE MODULE (INCLUDES /--MAINTENANCE HIGH PRESSURE CORE, UNIT COMPRESSOR REAR FRAME

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COMBUSTOR PRESSURE TURBINE /00 ..

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HIGH-PRESSURE ACCESSORY TURBINE DRIVE fAN AOTOR, STATOR AND FAN ROTOR 1ND FAN FRAME STATOR Figure 9. CF6-BO Engine.

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, , Figures 11 and 12. The fuel system arrangement i: typical of General Electric !.arge commercial engines. The CF6-BO engines (B-767 and A-3l0) include fuel cooling of the integrated drive generator (IDG). Air cooling is provided as an alternative heat sink for fuel cooling of the IDG oil. The arrangement is shown in Figure 13.

For this study, the engine fuel system components of particular interest t , are the fuel pump and fuel nozzles. The gear-type pump is shown in Figure 14. Although, as mentioned, the CF6-S0X was used for the study, it was decided to use CF6-S0 fuel nozzles in the Baseline model. This was done because good knowledge of field service performance is presently available for the -50 nozzle thus permitting a more accurate assessment of fuel temperature effe~ts. The baseline fuel nozzle (CF6-S0) and manifold arrangement is shown in Figures 15 and 16.

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Fuel System Schematic. i

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DEBRIS SeRF-EN ~ HIGH-PR[SS FUEl/OIl PRESSURIZING SHUI-O/f HlAr ()C~H PIlMP AND DRAIN VALVE VALV( r---'" --,

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Figure 12. Fuel System Schematic .

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i;:i;j fuel ~-. /, Fuel/Oil l":?~j 1 B

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Fan . '

Discharge • Air

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I Integrated ~ Drive Oil Generator J ,: 1 N N

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'j I~ Overboard Air

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• During Takeoff, Climb and Cruise ~ Air Valve Closed ....

.

IDG Oil Heat Goes to Fuel 0.35% .o.SFC Improvement at Cruise • During Idle Descent - Air Valve Open lOG Oil Heat Goes to Air • Used on CF6-80A For 8-767 and A-310 Figure 13. lOG oil Cooling Syste~.

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PORT fUEL HIGH.IHSUIE :_ I INLET IEUEFVALVE , fOlT JET EDUCTOR

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Figure 14, Fuel Pump, ~.

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VALVE ~r, HEAT SHIELD (SHOWN .. lilFWll l'illlWIliI PRIMARY

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PASSAGE J:ln:.;~'1 1:-11 ~ mr

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. , , flOW ONLY NOZZlE TIP ASSEMBL Y ~

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Figure 15. CF6-50 Fuel Nozzles (Pressure Automizi.ng>'

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SHROUDED FUEL NOZZLE FEEDER TUBE INSTALLATION

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FUEL SHROUD MANIFOLD ASSEMBLY PYLON TUBE ASSEMBLY .. (

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LOWER PYLON FIRESEAl ,.

FUEL SUPPLY ruBE fUEL FLOWMETER Figure 16. CF6-50 Fuel Nozzles with Valves on Casing.

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.1 4.0 ADVANCED SYSTEMS

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4.1 INTRODUCTION ::' I The system arrangements and component selections for the advanced systems are keyed to the issue of broadened-propert~ fuels. For the study, three fuel properties were of primary concern:

•

l. Freezing Point - tank heating consideration 2. Thermal Stability - fuel nozzle consideration 3.

Low LubriCity - fuel pump consideration.

The results of the study provide comparisons between three Advanced Systems (A, B and C) and the Baseline System. FiguL'e 17 lists the fuel properties and the approach or method for dealing with these properties. In all cases, tank heating is used to provide fuel freezing protection. Low lubricity problems would be avoided by either the use of low lubricity tolerant materials or the use of centrifugal pumps. Thermal stability problems are addressed by improvements in the fuel nozzle design and by reducing the fuel temperature supplied to the nozzles.

Figures 18 through ~l show, using a common fo~at, the differences between the Baseline and the Advanced Systems. The major features of each system are indicated by these figures. Note that these syste~ as fo~ulated in the study computer models start with the aircraft auxiliary and No.2 main tank. System weight comparisons are shown in Figure 22.

4.2 ADVANCED SYSTEM A In this system, a fuel return line is connected from the engine fuel I.

control to the outboard fuel tank of the c.orNsponding engine. Normally all control bypass flow returns to the tank. This pe~its a portion of the pump

L

heat and engine lube system heat to be transferred to the fuel tanks.

Explicitly, heat from these sources is transferred to the tanks or to the PRECI,WING PAGE BLANK N(r FILl\IED

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Ady,nsed ~'IIl!ll1I (!"Il:~Ql"EU- aOK Fuel Prop.!!!1Y Ba!\ellne APJImGb A Jl .&.

Fuel Freezing Po'flB • Tank If.ollng • None • New • Ne. • Now Jel-O + 5· F : -53 F Spec • 00011 Pump Fuel • WeIer-Roel,.-Loop.

• Gear Pump Fuel

IBrp." Roclreu.aUon. Recirculation

Jal-A + 5° F = _35 F • Deeed Atr Hoat 10 Tank

SludV lImli + So F : - 12 F • lube .nd Pump He.1 • lOG H .. llo Tank To Tank Lub,lclll/ e Conventional • Fuel Pump • New • Ne. • New N Cu"enl : High Comp.IIbIIIIV • LOlii'tubrlclty • Centrifugal • Low Lubriclly (Xl Siudy L1mll : Low G •• r Uaterl" Gear .lale,i';!

Thermal Slablllly (Mantlald) • Fu.IN ........ nd • Conventional • Ne. • New • Na.

Avg_ Jol-A : 300 F Mal V.I.I. CompaUblll1y

• High - ap • Low- ap • Low- ap • Low AP

Siudy L1mll : 255 Fill •• • Vlllle. on Ca.lng • ValveD on eating • Romole stngle VU'oie • Lower FUIlar • VoIvl. on Manllold • Low fuel.6T • Low Fu.ILlT • Low FUll AT Figure 17 _ Relationship of Advanced Systems to Fuel Properties_

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·':" l-:-:l ~ ~rl- .' ~1 I<'-~<' Air To i} ...

ECS and :.i1.~ ...

Ant~-Ice ..

Aux. or No.2 , .

Tank Hoi "

fan Air ... 1 Air ~compressor

Cooling H!{ Bleed

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Outboard ., J

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Recirculation On Casing

Baseline OG10-30 and CF6-80X. I

Figure 18,

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t: lowap

I Nozzles ~

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low-lubricity Tolerant Casing Gear PLUmp Figure ]9. Advanced System "A".

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Air 110 .. "~ ECS and 'i:; A:.Jx. lOr" NIO. 2 ?, Anti-Dee Tank ,.. Hot Fan AI ..

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fuel I Outboard Tank IDG ~ • I: i Main Fue~ Aur Tank Oil HX HX W ~' >-' ,~ ..

low-Lubricity LlUlbe

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k LowAP Tolerant Heat I--' NozzOes Gear Pump

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Figure 20. Advanced System "B" .

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~ Air To AUK or No.2 ECS and Tank AntD-lce .--1.-.., Air Hot HX Compressor .. I Fuel!r Water and Antifreeze Bleed Outboard HX ~ Recirculation loop Tank

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Main Tank -.- W N Starting Pump low OBI .-- ~P

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Nozzle :1

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FlUlel H Fuel I .. ~ --./

HX I IMECI-I -~

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Single ~ I ./1 CeUlltll'ofuga, Remote ~ Pump Valvi> ~ I Figure 21. Advanced System "e".

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~;; Bleed Fan Fuel Air Bleed WHRS Pump- PrecoQl~er Valve loop- Total

---

Base~llI1e

J 39/11.7

62/28.1 14/6.4 115/52.2 , , System A 39/17.7 62/28.1 14/6.4 115/52.2 System B.

39111.7 62/28.1 14/6.4 115/52.2 w w ) System C 29113.2 157171.2 186/84.4 ~ IncludulI1g j Start Pump , Note: Th~se Weight Penalties Increased by 25% in Model to , Account for Additional Engine Mount Weight

Figure 22. Weight Tradeoffs (Dry Weight, Pounds/Kilograms), I

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engine combustor in proportion to the ratio of tank return flow to engine burn flow (metered flow). In certain situations such as ,tank temperature exceeding limits or needing additional fuel icing protection, it becomes necessary to shut off tank return flow. This may be done by cockpit- controlled operation of e valve which causes control bypass flow to return to engine pump interstage (in the normal manner). The reason for suggesting tank return as a normal operational mode is to gain the added benefit of ~educed fuel temperature rise in the engine system. This occurs because the fuel temperature rise associated with pumps and lube heat are now (with tank return) chargeable to tot~l pump flow (gear flow) instead of only metered flow. This reduced fuel temperature is advantageous for cooling other systems and operating at lower nozzle temperatures. Fuel cooling of auxiliary systems I .j (such as the 100) results in better sfc:: by reduction of air cooling (see Figure 13).

System A uses low AP (air atomizing) nozzles with the divider and check

I

valves mounted on the engine casing. This low AP nozzle is of a design similar to the nozzles on the General Electric TF34, T700 and GFH56 engines.

These air atomizing designs are shown in Figure 23. From a thermal stability standpoint, the advantages of these nozzles are: (1) larger injector ports, (2) more air cooling of the nozzle tip, and (3) lower nozzle AP resulting in

I

lower pump pressure rise (less pump heut input). Reduction in-nozzle AP from the Baseline (high AP CF6-S0 type) is shown in Figure 24.

__ I For low lubricity considerations, system A uses pump gears made from a carbide matrix material. The gears are dimensionally the same as standard tool steel or nitraloy gears. various carbide-matrix gears have been and continue to be evaluated by pump manufacturers. Titanium carbide in an iron matrix is one possibi.lity, Tooth bending strength is imparted by the iron.

Titanium carbide is hard and self-lubricating because of its hexagonal crystal structure.

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4.3 ADVANCED SYSTEM B

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Advanced System B incorporates a fuel tank return line which receives

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heat input from the generator (lOG) oil system. The engine fuel pump centrifugal boost stage is oversized to provide additional flow capacity. In effect, the arrangement simply moves the lOG fuel-oil heat exchanger from its normal location (downstream of the fuel control) to the tank return line. The air-oil heat exchanger provides cooling if tank temperature exceeds limits.

, .

The same low 4P fuel nozzles are used as with System A except that the divider and check valves are located off the engine casing. Valve locations are shown in Figure. 25. Secondary flow nozzles are used only at power , .

settings above idle. If the check valves for these nozzles are located on the engine casing, the comhin,tion of static fuel and heat influx from the casing [315.6' C (600' F) at valve closure during power reduction] could accelerate

coking and affect valve freedom of motion. With these seconda~y-only valves I

i on the manifold, fuel flOWing in the manifold provides a heat sink ,-0 cool the valve continuously during engine operation. In addition, the manifold location more effectively isolates the valves from the hot engine easing.

, •• System B uses low-lubricity tolerant pump gear materials the same as System A.

4.4 ADVANCED SYSTEM C Advanced System C represents the greatest change from the conventional design approach. Figure 26 shows a more complete description of the Waste Heat Recovery System (WHRS) used to heat the fuel tanks. The system uses engine compressor bleed air as the source of neat for tank fuel freezing pro~ection. Normally air is bled from the engine for wing and engine cowl

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anti-icing and for cabin environmental control. Before the compressor bleed air can be routed away from the engine (to the wings or ECS) it must be i i .~ cooled. Engine fan bleed air is typically used to cool engine compressor bleed air. A large air-to-air h~at exchanger (precooler) is used for this J: ..

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Valve on Manifold for Heat Sank at Low Power (System 8) J Fuel Supply Figure 25. Nozzle Divider Valve tocations.

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Figure 26. Advanced System C.

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purpose. The arrangement can be seen ~n Figures 8 and 10. In the interest of reduced fuel consumption (lower sfc) the trend is towat'd less engine bleed extraction for the ECS by means of cabin air recirculation eystems. Studies are also in proc:,ess to eliminate fan bleed (and the precooler) by using selective bleed ports on the engine. Ideally. the bleed air would be taken from the stage of the engine that most efficiently and effectively provides the desired pressure (cabin pressurization needs) and temperature (cabin cooling or heating). If tank fuel heating were not an issue (fuel freezing point remained low). these multi-port bleed systems and O~;ler adl."anced concepts may be preferable to a waste Heat Recovery System. It should be pointed out. however. that a workable system has yet to be developed which meets these objectives (elimination of fan bleed and precooler).

Given the situation where both sfc and fuel freezing point are

I

significant issues. the WHRS is of particular interest. In this system. as I shown in Figure 26. the precooler is eliminated. Either (or both) engine fuel and tank fuel provide the heat sink for compressor bleed air cooling. There are numerable ways such a system can be put together. For this study. an arrangement was chosen which provides a high degree of design flexibility.

The jet engine is always viewed as a multi-application product. Consequently.

installation differences between different aircraft is a prime consideration.

Bleed air lines, are typically large in diameter and not routed in close ~, ; proximity to the engine fuel system. Consequently. a liquid transport system I ' "!:../o: .. "_ is desired to transfer heat from the engine bleed location to the engine fuel. The desirability of a transport system (as opposed to direct heat

l'

transfer) is reinforced by aircraft tank heating and cabin air conditioning considerations. Routing additional fuel lines to and from the tanks is to be avoided if possible. The possibility of fuel leakage into the cabin air system is also undesirable. Both of these sa~ety issues are addressed by use of a liquid transport loop.

'(' I, The system incorporates a water/anti-freeze heat transfer loop. The anti-freeze may be Dowfrost which is used in aircraft food chillers and is \

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non-toxic in the event of a leak to cabin air. An electric motor pump

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circulates the water between heat exchangers located at engine compressor

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bleed ports, engine fuel and tank fuel.. Water loop bypass valves automatically control the distribution of heat to either engine or tank fuel

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and also control the temperature of the bleed air. A fuel recirculation arrangement located near the fuel tanks provides heat transfer between tank fuel and the water. The water system is statically pressurized by engine

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compressor discharge air (COP) so as to prevent water boiling at low ambient pressures.

Advanced System C uses a single remotely located ruel nozzle flow divider valve. The arrangement is sho~~ in Figure 27. The valve would be designed to distribute fiow to the nozzles (injectors) as a function of upstream supply pressure. This valve would be thermally isolated from the hot casing and would operate at a higher force margin than the typical nozzle divider/check valve. Hence, from e thermal staoility point of view, the arrangement is

I

attractive. Low liP nozzles are also used. A disadvantage of the arrangement i~ ';he need for as many as 30 separat~ fuel delivery tubes leading I, from the valve to the nozzles.

System C uses a high speed centrifugal fuel pump and a throttling-type fuel control system. This is similar to t~e afterburner syste~_on military engines. A separate low-spee,d boost pump is used to pressurize the inlet to f! the high pressure pump. The arrangement is shown in Figure 28. Engine starting is accomplished by use of a separate slectric-motor-d.'iven centrifugal pump. The entire arrangement is conventional from a component standpoint and similar to existing centrifugal systems used on jet engines (Concorde SST, for example). The particular advantages of a centrifugal pump ::\ system are inherent low-lubricity toterance and low fuel. tel1ll'erature rise at idle flow rates. The fuel control ca~ be either hydromechanical or electronic. The throttling valve (see Figure 28) reduces pump discharge pressure to the value needed to hold a constant liP across the metering valve. For the study these valves were assumed to be a flat-plate shear type in order to increase tolerance to low-lubricity fuels. Flat plate (shear valves) can be easily made from hard carbide materials incorporating

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self-lubricatins characteristics similar to low-lubricity tolerant gear pumps. Reference 2 discusses these valves in more detail.

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5.0 FUEL PROPERTIES 5 • 1 FUli!L PROPERTY CONTROL

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Fi~ure 29 shows how the properties of aviation ~rade fuel are controlled. In the past there has been a wide acceptable crude selection in · .

the form of petroleum. This wide availability has been coupled with a relatively low airline fuel demand. Selection of the crude source has • permitted the desired end result in terms of properties which include aromatics, thermal stability, heating value and viscosity. Ordinary strai~ht distillation processes at the refinery lead to control of boilill~ .ran~e with corresponding control of freezing point and flash point. Airline demand for kerosene-type fuels such as Jet-A has prov;.ded a ~enerally good market , distribution of the total available crude product. Consequently, market cost , , factors (competin~ demand for end product) and actual cost fac.tors (cost .to

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, produce aviation spec fuel) have been in harmony .

~ In the future, it can be expected that t:.:re will be a ~eneral reduction in the total availability of acceptable crude sources (petroleum) as compared , , I.

to the past situation. This trend has been evident for several ye~.s and can 0' bE-: '.loted by the heavier, hi~her aromatic fuels such as those derived from the

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Alaskan Slop~. Fi~ure 30 shows generally what can be expected~n the future.

· .

It is irrelevant to the study as to when this situation may develop. An 1. aviat.ion fuel property iSSllri may emer~e under world economic conditions that reflect the rate of ~rowth (increase in GNP) which has ~enerally existed

l' durin~ the past three decades. Coupled wit.h this increase in economic ~rowth

may also be an increase in airline travel as a percenta~e of the total demand on transportation ener~y. These conditions (1) reduced selective crude sources, (2) economic ~rowth, and (3) increased airline demand for fuel may lead to a situation where fuel properties have a more deciding influence on fuel cost.

Over the spect~lm of potential crude sources there is a ~eneral trend away from property control by crude selection. The result is a need for BLANK NOT. FILMED PRECEDING pAGE

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Figure 29. Aviation Fuel - Past Situation.

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Aviation Fuel Future Situation Increased Aviation Supply at Present Spec Limits .

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1,("-- • additional and more complex processing in order to· arrive at the desired end result. For example, low lubricity is not presently a significant property issue. Severe hydrotreating removes impurities and hence the natural lubricity of these impurities. At present, however, there are few modern refineries using severe hydrotreating. In cases where aircraft fuel pump failures have occurred because of ;1 lack of fuel lubricity, the problem has been solved by additives such as corrosion inhibitors. Eventually, actual cost factors associated with the production of spec-grade aviation fuel must increase. This is because the aviation grade fuel will no longer be a simple matter of crude selection and simple distillation processing. Adding to an iucrease in actual cost may be an increase in market cost. This comes a~out ,~ ; as the demand for kerosene fuel increases. The higher fuel efficiency of diesel engines in automobiles leads to their de$irability. In a free market environment, both airlines and automobiles may be competing for essentially the same kerosene product. This would of course be a dramatic change from the past harmonious relationship between kerosene demand and gasoline demand.

Enough is known about future fuels to r~cognize that a choice may 1 I ultimately be made between rigidly defined aviation fuel properties and - !

broadening of these properties. Table 1 shows an assessment of aviation fuel change which might occur under assumed conditions of ~orld economic property Table 2 shows present specification limits for Jet-A anQ Jet-B growth.

fuels. Also shown are limits for the study fuel. These study fuel property values are given here only as a matter of reference. They assume an increase of about SO percent in aviation fuel yield from the crude sources listed in Table lover the next 15 to 20 years. It is irrelevant to the results of this study whether fuel properties change to this degree or only slightly from present-day values. The systems evaluated by this study afford various degrees of property change. Airline operational procedures also are a factor in equating the performance of a system to its fuel property compatibility.

If one assumes that fuel properties will change or that lowering of relative fuel cost would make a property change desirable, the next question is how to accommodate the change from the standpoint of the airline business. . , i , ..

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r~~"': I)'; ! -/~, r, .'l" TABLE 1. ESTIMATED FUEL PROCESSING REQUIREMENTS ,.

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[ower Freezing Thernlill Flash Aromattc Heating !iOl!rce Lubrtctty Point StablHty Point VhcosUv Content Va~ue Petro]eu. CrUdes (average) 1980 - 1990 U 0 0 0 1 0 1990 - 2000 2 1 0 1 2 1 2 • Synthetic Crudes .,.

) '" - Tar Sands (1980 - 3 1 0 1 3 2 3 ) - Shale (1985 - 2 2 0 2 2 3 2 - CoaU (1990 - 3 2 .. 3 ..

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» ----- -- .. ------ .. ~ --- -- ------- ----- -- - - Processing Code: o - Similar to 1970 Crude - Meets Currenl Jet-A SpeCification.

1 - Worse Than 197101, - But ~xpected 10 Meel Specification Without AddWc'!"!aO Processing I 2 - AddilOollial Processing ProbabOy Needed to Meet Specification with Adequate AvaOlabilOty 3/4 - Significant Additfionai Processing Required 10 Meet Specifocation

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" TABLE 2. FUEL PROPERTY SPECIFICATION LIMITS ~~.

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',~ lower Heating freezing

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Thermal Flash Viscosity Aromatics Valve Point Stability Point ~ at -20° C Content BTU/lb ° F/(OC) °F/COC) °F/(OC) CS Vol. % (kJ/kg) Jet-A -40 Max 473 Min 100 Min 8.0 Max 25 Max 18,400 Min (-40) (245) (38) (42,800) V1 Jet-B -58 Max 473 Min N/A N/A 25 Max 18,400 Min (-50) (245) (42~800) Study Fuel* -17 Max 428 Min 80 Min 10.5 Max 35 Max 18,300 Min (-27) (220) (27) (42,570) * Estimated 50% Increase in Yield

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Table: 3 Hsts several options available to the aircraft user. Ideally, fuel

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properties and the costs of meeting these property limits is precisely matched against airline profitability and passenger safety. Otherwise, money is being lost, or accidents are too frequent. Assuming this ideal relationship to now exist, each of the fllel properties listed in Table 3 will (if changed) cause a ,_a , ~ " need for a corresponding change in order to preserve this idealized balance.

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Undpr these assumptions, flight crew action is inappropriate for any property change. Increased maintenance can only be applied to the economic side of the issue (not the safety consideration) if the result is a direct change to aircraft performance or reliability. Totat is, if an increase in maintenance is necessary to accommodate satisfactory aircraft performance or reliability, and therefore not jeopardize safety, only economics are aff.ected. Operational , 't ,.

planning influences both economics and safety if freezing point is increased or heating value is decreased.

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Freezing point, therma.i. stability and lubricity are the properties from

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Table 3 ~ich are most meaningfully ad~ressed from an aircraft/engine fuel L system standpoint. The other properties are involved more with the engine combustor (viscosity, aromatic content, and heating value) or fuel handling , safety (flashpoint). In the study, freezing point was ~onsidered from two ., points of view: (1) operational consideration and (2) design consideration.

!

Thermal stability was considered both from a design and maintenance standpoint. Lubricity "'as considered only from a design standpoint.

( 5.2 ~~~IONAL CONSIDERATION OF fREEZING POIN1 Figure 31 shows how fuel fre~zing point is addressed by the airline. The airline captain responsible for the safety of the aircraft does not know what the exact freezing point* of the fuel is during any varticular flight. He knows, however, the fuel type and its specification limits. He also knows and I L has available the aircraft flight manual which describes any limitations associated with cold fuel operation. He assumes based on his flight engineer's observation of fuel tank temperature (ucaally one measure of bulk

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*Temperature at which last waxy solids melt when the fuel is ~eated.

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Freezing Point x

" Thermal Stability X X U.

N Flash Point X _I Viscosity X . , X i X Aromatic Content Heating Value X X , X Lubricity , '--- \;.'V * ' '" '",' -, 1: .- .-';....;;.;""....;-'."."~". _____ ~~"-~~ ___ _

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CD Pre-Flight Planning (Airline Practice)

" , Fuel loading Fuel Selection • Reute

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• Time ef Year • Fuel Will • Jet-A

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i Stay Above • Experoence • Jet-A1 FP+ 5° F (3° C) , • Jet-B • Seasenai Adjusted Freezing Point V> W , , :

® In-Flight Control (Not Done)

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temperature in one tank) that all fuel is available except that which is consumed as the flight progresses. Typically. the aircratt flight manual will permit operation to the fuel specification maximum freezing point plus 3° C (5° F). With present day fuel properties. it would be a rare occurrence to reach this limit. Figure 32 goes further to show airline practice with respect to fuel freezing point. By using a grade of fuel (Jet-Al or Jet-B) with a lower freezing point. the airline is able to stay clear of potential fuel freezing problems.

In the future. this situation may change. Such change may be predicated on economic factors such as fuel cost relative to fuel properties. The following factors can be considered: • Increased probability of freezing point influence during certain flights.

• Meaningful cost trade-off for fuel selection.

• More advanced computer techniques available for en route weather forecasting and fuel temperature forecasting.

The factor involving en route forecasting was used in this study to predict what might happen in terms of fuel tank temperature. Figure 33 shows , .

. J , the approach to operational planning that an airline might use in the future.

For this study. the analytical approach was as follows: 1. A flight forecasting technique was developed in a manner which might be similar to that used by an airline. The objective of the forecast was to obtain fuel tank temperature.

2. The baseline aircraft and engine was flown by computer simulation through the forecast air temperature and route structure.

3. Each Advanced system was in turn· flown through the same forecast flight.

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r' ~j Federa~ Aviatnon Regulations (FAR's)

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1 i , Fuel Specoficatftons Max Freezong Peont I' en (FP) en Flight Limit Aircraft

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~ FP + 5° F (3° C) For Irnstrumentalion Err'rlrs and Temperature 1 Varna loons Figure 32.

Freezi.ng Point Current Operation Requirements.

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• Reule Altitudes , • Air Speeds City Paors Air Traffic ~ • AII' Temp Aircraft Centrel Weather

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• Fuel Required (Gly) loading Forecast • FlUgh~ Time \n '" I FueD Temp •

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Fuel Special Fuel Selection Availability Equipment fer lowes! ClOst Spec limits Figure 33. Operational Planning for Lowest Ftlel Cost.

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6.0 METHOD OF ANALYSIS AND FLIGHT DEFINITIONS I J ~ • 1 THBRHAL MODEL ..

A comput~r model was formulated to represent the baseline aircraft and engine. Usirtg common formulations as appropriate, additional models were prepared for the advanced systems. These models were extensions of the General Electric CF6-BO engine-family heat model used to predict engine fuel, i .

lubrication and generator oil temperatures. In the CF6-BO application, the validity of the model has been verified by engine ground and fligllt test data such as for the Boeing 7671General Electric CF6-BOA. The models follow thermodynamic 1st-Law expression of the energy continUity throughout the aircrsft snd engine fuel systems and thermodynamically interconnected oil al.', air systern'3.

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Figure 34 for the baseline model shows a simplified description of the model. Fuel from the aircraft auxiliary or number 2 tank (inboard tanks) is pumped to the main engine outboard tank. The temperatur& of these supply tanks changes as the flight progresses. Heat (Q) is added by the tank fuel

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transfer pump. In the outboard tank heat (QA) is transferred to air flowing over the wing. Fuel flows from the outboard tank overfill valve to the main

I . tank. Hore heat transfer occurs to air over the main tank wing-area. Engine

meter flow (WFE) is pump~d through a main tank boost pump to the engine. The i' \ mass of fuel in the main tank (H) decreases. In the engine, fuel is pumped ! .

through the engine lube system he~t e~r.hanger. Excess pump flow is throttled with corr~sponding temperature increase (Q). lOG heat is added to the metered

\

fuel and then i!lOre throttling Emd temperature rise occurs at the fuel , nozzles.

I Figures 35, 36 and 37 show the typical arrangement for fuel return to the aircraft tanks. In order to suppress the cooldown rate of the outboard tank, tank return is made to the outboard. This results essentially in a re~Lrculation loop from the main tank to the engine fuel system to the outboard tank then back to the main tank.

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, ; I I I I I No.2 II

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.-- Y-IDGr~

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E I ~ , WFE Figure 34. Energy Transfer - Base!ine,

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j F.igure 35. Energy Transfer - System A.

l

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HX E WFE Figure 37. Energy Transfer - System C.

lv __ "' ______ ~_..._..__~~~--, .............. ',. . ...

__ _liP C:O::S-=::

Figure 38 shows the basic thermodynamic considerations involved with the various fuel tanks. Fuel mass may be either constant or changing at dif.ferent times during the flight. Total thermal enl'!rgy (total BTU of thermal energy in tank) and specific energy (BTU per pound of fuel in tank) are ac~ounted for ill the model. Figure 39 shows the component energy exchanges which occur in the system. Total thermal energy of the flowing fuel is bookkept in the manner shown. Note that throttling results in an increase in total thermal energy since pressure energy is expended in the constant enthalpy process.

Figure 40 shows the heat transfer mechanism occurring in the tanks.

Minor effects such as radiation are not included. Fuel in the tank exchanges t.herrnal energy with air outside the wing by the process of forced convection in the wing boundary layer air film. Air film convection is predominant over other heat transfer modes in the tank (liquid to tank walls for example). The tank wall and the fuel bulk temperature are identical in the model. The air temperatu!'e if.volved with the heat transfer process is the wing boundary layer temperature which is the recove .... y temperature (T ). Note tllat recovery R temperature is lower than total temperature since stagnation occurs only at the leading edge of the wing. Air heat transfer properties are calcu'lated from the air film temperature which is a function of wall (metal) temperature and recovery temperature.

Table 4 lists the steps used to calculate the tank-to-air heat transfer rate. Fuel volume affects the heat transfer area by means of both wetted internal surface areas and conduction through the tank stringers.

The thermal model calculates temperatures, pressures, flows, heat transfer rates, and sic effects at one-minute intervals through~ut the flight. Tank t'3mpe~ature determination is crucial to the accuracy of the results since virtually all other systems cQnditions are influenced by the engine fuel inlet temperature. Figure 41 shows how the tank temperature is determined. Tank fuel temperature (T ) is the temperature of the bulk fuel at the end of each one-minute interval. Note that l.oookkeepitig of total :"j thermal energy (E ) and fuel mass (M ) form the basic thermodynamic 2 2 ,- statement of accountability.

._.

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,-- .- .......

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.. Fuel Mass Constant .. ~

.. Temp Increase/Decrease 'j .. ... QA (j) (j) -1 .. Total Energy Increase/Decrease · ~ I .. Specific Energy Increase/Decrease .. fuel Mass Decrease

(2) ~ = -W91.:J

.. Temp Constant -~ ... ~ W

<ID

w '" • Total Energy Decrease .. Specific Energy Constant .

.. Fuel Mass Increase T~vot .. Temp Increase/Decrease @

w

@ .. Total Energy Increase .J .. Specific Energy Increase/Decrease Figure 38. Basic Thermodynamic Consider"t;ons for Tanks.

~ . -.t .• ,_, ,,- t •• _.

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• E2 = E1 + HP- ~p (GPM)/1714 @ Throttling (Pressure Drop) , I • W2 = W1 L.J T W - ~ -:.~--:=-_--:-

.'

• P2 <P1 1 fI 2 • E2 = E1 + ~p (GPM)/1714 Figure 39. Basic Thermodynamic Considerations for Components.

E = Total Thermal Energy (BTU)

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v = Ale Speed

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r QA-'~

0-.

~I en L

-I

TW = TFl

7/ZTJ/fl!2Z?J

, ' ':

Aftr Film. T A* = '(TW. TR)

,- ....... u -.-'~ Recovery Temperature T R = (1 + 0.18 Mp2)T Slatic Figure 40. Tank Heat Transfer to Air.

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~

.~ ________ .;._..;. __ " •.• "",-, ·"_.;':'.~b .. _ -~

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F: / r (~ ~ . ' TABLE 4. TANK HEAT TRANSFER TO AIR 3: d) Define:

:1

• Inboard or Outbclard Tank - for Tank Slenderness • 100% fUf}1 Volume - Tank Capacity ft • Wing Cord at Tank - for Air Reynolds Number (2) Find: '" '" • Actual fuel Volume (Level) - ft3 j • fuel Tank Wefted Area e function of % Vol..

Inboard/Outboard, A/C Size • Area Increas, for Stringer Contact With Unwetted Area ~ K1 • Air film Temperature and Air film Coefficient - U .J ~ (3) QAir = UAK1 (iF = TR)

j

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t! o. I j

--- . : :-'0'1 !

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1, 1, For One-Minute Interval Time (j) To Time ~ Total Energy in Tank = M (AT + BT2/2)

.j

, < Energy Added/Removed By Flow = W (AT + BT2/2)

Energy Added/Removed By Heat Transfer = Q

'" " Total Energy E2 = E1 - E [W(AT + BT2/2)] - Q Mass M2 = M1 - EW ,

Specific Energy E2 = E2/M2

: ~ Temperature T 2 = -A + '" A2 + 2BE' , J B Figure 41. Tank Temperature Determination.

r

~ _ ' .. " --/-'.-"1.. .. ... 1-,

, • '!

I 6.2 MISSION DEFINITIONS In order to evaluate the differences between the baseline and advanced systems, it was necessary to simulate four flight missions of the DC-lO-30 aircraft. These missions were: 1. A one-day-per-year long-range cold environment flight originating from a cold airport Which results in the coldest fuel in the aircraft wing tanks d~ring cruise. The proximity of the temperature of tank fuel to fuel fr.eezing point shows the effect of this fuel J ; property on aircraft operational capability.

2. A one-day-per-year short-range hot environment flight originating from a hot airport which results in the hottest fuel in the aircraft wing tanks at the initiation of iile descent. This results in the hottest fuel normally in the engine fuel and lubrication systems.

Fuel thermal stability would influence the effect of these hot fuel temperatures on fue.l injector coking.

3. A nominal (50 percent probability) flight at nominal environmental and airport temperatures. This flight serves as the primary basis for assessing economic factors and sfc when system Qnd component changes are considered for resolution of the problems prespnted by fuel properties.

4. An emergency condition with low fuel reserves. This analysis was needed to assess the likelihood of excessive tank temperature caused by the tank heating system.

All of the mission analysis work was performed using computer models.

Figure 42 shows the anallsis flowchart. Tables 5 and 6 list the procedures.

e._ .~-" ..• i'

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Mission Definition Run Ellglne Cycle Deck Run DC10-30 Thermo! Model

I

j

• Choose Departure Airporto lor • Installed Thrust and SfC • Aircrall and Engine Fuel 8'1d ~ Cold, Nominal and Hal MI.olon8 • Engine Cvcle Dala: Fuel Lube TempI. Flow., Pre~~ureB - Boeing "Airport Temperature." F!ow, Speod., Engine Aero • Fuel Propertiel

r+ 'i

Manual Dala, Etc.

• aSFC - Performence and Conlrol • Chooue City Pairs for DC10-30 Operalion Considering Frequency Sys!. Eng'g Cf6-80X

I

and Length 01 Flights Cycle neck , - Commercial Forecsollng Roule

-

!.

Siructure Analysis Program Companenl Economics • Provide aT Ambient irom ISA and Headwinds at Altitude lor Season Analysis and Probability Englnel Alrcrall IntegraUon .. Acqulsilion - Boeing "Wlndtemp" Program :'f 0\ COBts • Reliability

[ '"

• Flight Prollle: Altitude, Mach No., • Airline • Maintenance • aTlme, aFuelBurn, aDlslance Thrusi Ral/ng ~l Operating l' Between Proille Polnls, Power Sell, • Call.

- Douglas Alrcrall ~.:. , COltl Fuel Loads/Reserves (IFR) Perfo; ;nance Manual

i

- "CAMAL" Program Using ..

I- - DAC Manual lor Alroral!

Welghis, Geomeirv, Aerodynamics, and Fllghi Prolile • Collata Dal~ and Distribute MIBllon Points In One Mlnule tnier'sl.

lor Thermal Model I!!pul { I , ,

j

Figure 42. Analysis Flowchart.

L

~~~~ ..

_ Hi Or '!?e"

TABLE 5. DEFINING FLIGHTS 1 Define Fuel loading Temperature • From GE/TBC CF6-80/767 Requirements

• One-Day-Per-Year Hot = 1010 : F (38° e), Nominal = 60° F (16° C)

i

One-Day-Per-Year Cold = 0° f (-18° C)

2 List Possible Departure Airports Using TBC "Airport Temperatures" • Cold 160 listings =20 io -78° F (-29 to -SiD C) Winter lows " a • Nominal 50 listings Most frequent City-Pairs • Hot 100 listings >115° F (46° C) Summer Highs 3 list DC10-30 Airline Departures Using" Airline Route Structure Analysis Program" • August 1981 Route Data I I • Screen For Flight length and Destination Airport • list Possible City-Pairs

j

!

4 Cross-Check Departure Airports with City-Pairs Ii Ust Possible Flights "- ~ .. '- ~ .... -- ;! :.

"'~, ...

~",'C"l ~;10.·"':i :c...-l f"-l :' '~"I TABLE 5. DEFINING FLIGIITS (CONCLUDED) 5 Define Cruise Conditions Based on DC10-30 Flight Manual • AUtitudes ana Air Speeds ., Define Altitude Tempeiatures and Head Winds Using TOe "Windtemp" Program • Average Values at Cruise Altitudes .- ....

• One-Day-Per-Year and Nominal Probabilities

-

,r:-., \ ;i :~, Selected Flights h , • Actual DC10-30 Scheduled Flight • Actual Routing • Probable Fuel l-oading Temp • , , • Departure Airpc.Plt Temp Consistent with Fuel Temp I ~ J ·1 , • Actual Flight Profile • Probable Air Temp and Head Winds

I

.

b

-= 3RT . w'" h. _.;...;.;..;.._-',"·OL", _____ ~_"_'__"' __

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TABLE 6. OBTAINING CYCLE DATA.

r ,.

" Known: " • Cruising Altitude ) " j,

• Cruising Altitude TemperaUures •

Jl 'i .. Cruise Air Speeds

• sIm Air Miles

..., N

<D Defone Remainder of flight Profile From DC10-30 Flight Manual

s Taxi, Ta~eoff, Climb, Step-Climb, Descent, Approach, Land, Tax!

® Define Altitude Temperature Profile Based on Departure, Cruise, Destination

, Temperatures

@ Find Fuel Burn From Commercial Aircraft Mission Analys~s (CAMAL; Program

• @ Dafone TakeoU Fuel load and Payload Based on Fuel Burn, FAR Reserves, J and Landing Weight Limits ~

l

® RUlli Engine Cycle Deck - Siore in Memory Bank lor Thermal Model Access ·4

l

, , .

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, , ~~l ~ ,.. iU :aiM . ..:..- --J"'.':iii"" ..... ,:....

6.2.1 AIRPORT TEMPERATURES The fi~st step was to identify wo~ld-wide ai~o~ts f~om which the : .

airc~aft might be depa~ting. Boeing "Ai~ort Tempe~atu~es" (Reference 5) was used to establish a list of depa~ture ai~orts which would have a likelihood , ' of cold. nomir~!.. or hot fuel loading temperatures to be used in the study.

Figure 43 shows. for example. the surface temperature at Helsinki. Finland.

'J Using coldest zero-probability surface temperatu~es. more than 160 cold-day departl,~e ai~o~ts were identified. Helsinki has a -33.go C (-2go ~) Feb~ary .tow temperature. Other locations had temperatures ranging from -28.go C (-20° F) to -61.1° C (-78° F) (Upper Canada during January). Surface temperatu~es at SO% probability were also listed to arrive at a likely choice of departure ai~orts for cold fll.ghts with cold on-loaded fuel. Similar departure ai~ort lists were const~cted for raminal and hot departure ai~orts using both zero and SC-~. probabilities, The nominal flight with 50 lieted ai~orts considered the m~. frequent fligh~ routes. For. the hot flight. Phoenix to Las Vegas was chosen. Phoenix hae a 47.8° C (llSO F) high tempe~ature. Eighteen DC-lO-30 ai~orts have highs in excess of 43.3° C (110" F). Over 100 world-wide airpor-ts have highs above 46,1° C (115° F).

6.2.2 COMMERCIAL FORECASTING R0UTE STRUCTURE ANALYSIS (CFRSA) The next step was to dete~;';cine fo~ all the departure ai~orts which ones operate DC-IO-30 aircraft. Figure 44 shows ai~line route guide data for August 1. 1981, for flights originating from Gatwick airyort, London.

Ensland. This offi~ial airline operation data is updated in t~~ GE CFRSA prog~am every th~ee months.

Gatwick was one of the departu~e ai~orts on the list for nominal flights.

DC-IO-30 flights were identified using CFRSA and the departure ai~ort listings for cold. !Iominal and hot su~face air temperatures. Screening of airport city-pairs was also based on fli.ght length. As shown in Figure 44, CFRSA uses great circle statute miles.

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......•..•......................•....•.................. -----_._._--_._--.-.--------

__________ fl ____ • ____ • ___ • _____ •• ____ • _____________ ~ ______ • ____ w ••• ___________ • ___ .__ OJ5'

OUT iTOP TYPI OF AIRCAA'T StoP ALN GRIO OUT SlOP OUT .... ,tOP

illlLl ... __ ....... ____ TOTaL DIll

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Figure 44. Airline Route Structure Analysis Program.

_nuuc.

, • • 1 6.2.3 ENVIRONMENTAL ANAL~SIS The Boeing ''Windtemp'' program (Reference 6) was then used to complete the screening on the basis of cruising altitude environmental temperature and geographical routing. Figure 4S shows the list of cold flight DC-IO-30 aircraft city-pairs which were analyzed using Wicdtemp. All results in the Wind temp computer program are computed in terms of nominal or S~ probabilities based on accumulation of world-wide weather reporting. For three winter months tl,u average temperatures at 39,000 feet over the great circle route b~tween Helsinki and Seattle is ISA (International Standard o 0 o 0 0

Atmosphere) minus 3.S C ~ 7· F), or -56.1 C - 3.S C = _60· C (-69 F -7 F

=

,{ , F) .

-76" The standard deviation from this value based on original data is

±

5· C (g' F). Assuming normal statist.ical distribution, two standard deviations would be appropriat~ for one-day-per-year. Hence, the average route temperature at 39K is -60· C _2(5° C) = _70· C (_76· F -2(9° F) = -94 F).

It is important to note that this is the altitude "average" static temperature. This value was arrived at in Windtemp using 200 nautical mile temperature nodes between Helinski and Seattle. At 39K, the outside air temperature will vary from this average, both warmer and colder.

Numerous factors affect tank cooldown ~ate: Ambient temperature Mach number Initial fuel loading temperature Fuel tank loeation Fuel transfer rate and schedules Tank size, shape, and internal design Fuel quantity Wing shape Altitude step changes Direction of flight

• Cold Flights

• Possible DC10-30 Coty-Pairs

WINDS AND TEMPS INPUT Ott I 1 20 5 18 10 19 10 6 6

---------2

"35000'- ---._- ----.---- 450--~-' - -- ---2-:-- 6

-0 - 00000020 BGR lGW 00000' .0 DEN LGW 00000.20 ANC NRT 00000210

'iNC-;m---------------------

--------- 00000220-- ANC CPH 00000230 ANC BRU 00000240 00000250

U··~~:~--i.~-i------ ----- - - -----

00000310 SVO lHR 00000320 SVO TAS 00000330

svo KDL

00000340 ---- - ---" --.-.----- ... - -----00000410---· TAS-S\io --. - .. -.

lAS FRA 00000420 KBl FRI\ 00000510 KBL COG 00000520 "HE[-:VUl-"-- -"-00000610- -- ....

Seiected

.... HEl JFK 00000620

HEl SEA oooooa30

City-Pair

AMS 00000710 .- ---------- - .-. -00000810--'--

AMS-

HeR-Sea

YUL ZRH 00000820 VUL MXP 00000830 YUL HEL 00000840 VU["-FCO---------- --------- ------------------- -- ---00000850-- 'iwG 'iVR 000009.0 'l''''IC; AMS 00000920 'HZ 'I've 0000'0'0 'iVl -YV~ 0000i020'-- vYZ AMS 0000'030 'IV i FRA 0000t040 YYZ MXP 00001050 - -- -'--- - .------------. ----0000.060---- YVZ-FCO'-- - VYl HNt 00001070 YVR YWG 0000 •• 10 YVR VYZ 00001120

-------_._- -- --

---- --0000Ti30--"-- YVR H~rL YVR LIM 00001 '40 'I've yyz 0000.2.0 'I've J\M~ 0000'220 9' - --- --00010000---'- Figure 45. Windtemp Input.

r.:""\l l4}o J When these factors are considered along with the fact that the aircraft type and airline routing do not necessarily correspond to a worst case transport aircraft scenario, it is evident that results such as these are "typ ical," not "worst case." Iience in a "typical" sense, the objective became one of

'-

I comparative results, not absolute results. In other words, if it were possible to show by analysis that the fuel in the left ci'ltboard reserve tank

I

of the OC-IO-30 flying from Helinski to Seattle reached e~actly -40· C (-40· F) at exactly 2 hours after t~keoff exactly once each year, this result would depend on precise statistical analysis of all relevant factors and even then would not represent the one-day-per-year or worst case for the airline ,j industry. Consequently. a significant issue i,,~ the coldest fuel temperature I which can be reached based on the possibility of cold ambient exposure somewhere along the route. For this reason, the city-pair or route ave~~ge ambient temperature used to analy~e other aspects of commercial aircraft mission analysis was also a logical choice for the analysis of tank coolciown rate and tank temperature. 'I'his method was in fact the only one available " which would consider a large number of mission options and provide consistency in the method of ati:llysis.

'The Windtemp program also ,Jrovide,d route-average headwinds and tailwinds. Winds were computed in the same manner as temperatures using 200 nautical-mile node points along the route to find the st.atistical city-pair average wind component. Wind was then added to or subtracted from great circle distance to find the air miles used t.o determine flight time and fuel burn.

6.2.4 AIRCRAFT AND ENGINE ANALYSIS Before \:he thermal model could be ::-un, it was necessary to determine the exact flight profile and engine power settings and then obtain the engine cycle data. This involved consideration and analysis of a number of factors and the utili.~ation of the Commercial Aircraft Hi.ssion Analysis (CAMAL) program.

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Fuel loading was determined based en intercontinental or domestic IFR fuel reserve requirements and takeoff/landing gross weight limits. The DC-lO-30 flight manual and cruise conditions frem Windtemp (altitude, temperature, winds) were used to construct the flight profile. During descent or climb transients, engine power setting adjustments were calculated at 0,5 to 2.0 minute intervals. These points were later fitted to constant one-minute intervals used in the thermal model.

Tables 7 and 8 show the flight selection results. Figures 46 and 47 show geographical locations for these flights. With the exc'lption of the 'not fllght (Phoenix to Las Vegas), these are actual DC-10-30 ticket guide rQutes.

,I The hot flight was assumed to be representative of a short hop for a wide-body jet although the DC-10-30 is not routed for this pat'ticular flight.

p ~ 6.2.5 EMERGENCY COlllDITION WITH LOW FUEL RESERVES To consider this issue, it was "'r. .. ~'med that all of the systems would continue to transfer heat to the fuel tanks in spite of whatever tank temperature occurred. The result would indicate the need for diverting heat input from the aircraft tanks.

The flight scenario assumed that the DC-10-30 approaches Phoenix to land with 37.8° C (100' F) temperature fuel in all tanks. For this emergency, fuel 1--- reserves are below minimum but still adequate to abort the Phoenix Landing and divert to Las Vegas. This is well within the design/perfot~nce intent of the aircraft. Fuel has already been transferred from the outboard tanks to) the mains as the aircraft approaches Phoenix. There is only 400 pounds of fuel in the outboard tanks. All engines are drawing fuel from the mains and t'l.nk heatir.lg is being returned to the outboard and then to the mains as is necessary for the operation of the advanced system.

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-- -- -.---~ ,~,. -~ , .:Mo._' "&.nI. , TABLE 7.

SELECTED FLIGHTS.

FI!ght Nominal Cold Hoi Gaftwlck, london to Helsinki, FlnOand 80 PhoenIx, Arizona 00 Cnty Pair Kennedy, New York Sea810e, Washington lal Vegal, Nevada (LGW --- JFK) (tiEL --- Sea) (PHX --- LAS)

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Deila T Amb from ISA/Probablllay +2° F (+1° C)/50% -25° F (-14° C}I +34° F (+19° C)I 1 Day Per Year 1 Day Per Year Winds at 50% Probabllily (Knoll) -37 (-69 km/hr) -13 (-24) +3 (+6) <!Areal Circle iIllstallce (NM) 3005 (5570 km) 4140 (7670) 222 (411) Equivalent Stm AIr Distance (NM) 3275 (6070 km) 4263 (7900) 220 (408) Cruise Almudes(s) (fl) 35K' (16).7 kll1l) !liliial 35K' (10.7) InGMal 20K' (6.1 kin) Cruise ~Iep Climb fto Crulle Step Climb 10 39K' (11.9 kll1l) 39K' (11.9 kll1l) Reserve Fuel Requirements FAR International FAR Infternaft!onaO Max Fuel load Posslbie , WOIhoul Exceeding Max

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Landing Welg!liU Payload 60% 0' Max 60%0' Max

4111% of M:;~ I

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!," TABLE 8. FLIGHT TEMPERATURES.

;: Hot Co~d Nominal Flight PHX--WAS HEl--SEA lGW --JfK coty-Pair 20K' (6.1 km) 39K' (11.9 km) 39K' (1109 km) Max A~t

. !

July Annual Jan Month 0 _6 F (_21 C) 0 _69 F (_56 C) _69 F (_56 C) 0> ISA Temp ....

0 +4 F (+2 C) _1 F (_4 C) +20 F (+1 C) /J,T Amb (51Di% Prob) +30 F (+11 C) _18 F (_10 C) 20'{One Day Per Year) +28 F (_2 C) _94 F (_70° C) _67 F (_S5° C) ADt l'emp

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