A FUSELAGE I TANK STRUCTURE STUDY for ACTIVELYCOOLED HYPERSONIC CRUISE VEHICLES - SUMMARY LOAN COPY: Retul2bj &FWLTECHNICAL LIERARY C

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NASACONTRACTOR REPORT

A FUSELAGE I TANK STRUCTURE STUDY FOR ACTIVELYCOOLED HYPERSONIC CRUISE VEHICLES - SUMMARY LOAN COPY: RETul2bJ &FWLTECHNICAL LIERARY C. J. Pirrello, A. €3. Buker, KIRTLAND AFB, Ne und J. E. Stone

Prepared by MCDONNELLDOUGLAS CORPORATION St. Louis, Mo. 63166 for Langley Research Cezzter

NATIONALAERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. FEBRUARY 1976 I " TECH LIBRARY KAFB, NM

OObL498

Report No.1. Report 2. Government Accession No. 3. Recipient'sCatalog No. NASA CR-2651 ". .~ _" _" 4.

. . .. .~ "~- "" ... - - ~- t7. Author(s1 !8. Performing Organuation Report No. C. 3. Pirrello, A. H. Baker,and J. E. Stone I 10. Work Unit No. "~""_ "" - 9. Performing Organization Name and Address McDonnellDouglas Corporation 11. Contractor Grant No. P. 0. Box 516 St.Louis, MO 63166 NAS1-12995

__"__~___ " 12.Sponsoring Agency Name and Address ContractorReport

NationalAeronautics and Space Administration 14. SponsoringAgency Code Washington, D. C. 20546

"" .- .- ~~~ I 15.Supplementary Notes

FinalReport

"" ." " ~ " -" .- - .. -. . - r16. Abstract A detailedanalytical study wasmade toinvestigate the effects offuselage cross section (circular and elliptical)and the structural arrangement (integral and non-integral tanks) on theperformance of an actively cooled hypersonic cruise vehicle. The vehicle was a 200 passenger,liquid hydrogen fueled Mach 6 transportdesigned to meet a rangegoal of 9.26 Mm (5000 NM). A variety oftrade studies were conductedin the area of configuration arrangement, structuraldesign, and active cooling design in order to maximize the performance of each of threepoint design aircraft, circular wing-bodywith non-integral tanks, circular wing-body withintegral tanks and elliptical blended wing-body with integral tanks. Aircraft range and weightwere used as the basis for comparison. The resultant design and performance characteristics showed thatthe blended body integraltank aircraft weighed the least andhad the greatestrange capability, however, producibility and maintainability factors favor non-integral tank concepts .

" "_ .~ ." " ~" ," - - . - ~ "" - - ...... - I." . " . - 17. Key Words(Suggested byAuthor(s1) I 18.Distribution Statement Hypersonic Aircraft Unclassified - Unlimited ActivelyCooled Structures ThermalProtection SubjectCategory 39

I ." - .. " - ... -. ~ ~ - I ~~ L- 19. Security Classif. (ofthis report) 20. Security Classif. lof this pagel22. Price' 21. NO. of Pages U nc lassifie d Unclassified Unclassified 77 $4.75 1 ... ~"". "... . For sale by the National Technical InformationService, Springfield, Virginia 22161

FOREWORD

The basic purpose of this study was to evaluate the effects of fuselage crosssection (circular and elliptical) and structural arrangement (integral and non-integraltanks) on theperformance of actively cooled hypersonic cruisevehicles. The study was conducted inaccordance with the requirements and instructions of NASA RFP 1-08-4129 andMcDonnell Technical Proposal Report MDC A2510 withminor revisions mutually agreed uponby NASA and MCAIR. The study was conductedusing customary units for the principal measurements and calculations.Results were convertedto the International System of Units (S.1.) forthe final report. Detailedresults are givenin the following reports: NASA CR-132668 AircraftDesign Evaluation NASA CR-132669 ActiveCooling System Analysis NASA CR-132670 StructuralAnalysis. The primarycontributor to thecontents of this volumewas C. J. Pirrello. Assistance was providedby A. H. Bakerand J. E. Stone.

iii I . . TABLE OF CONTENTS

Section Page

FOREWORD ...... iii

LISTOFFIGURES ...... vi LIST OF ABBREVIATIONS AND SYMBOLS ...... ix

SUMMARY ...... 1

INTRODUCTION ...... 5

PERFORMANCE AND DESIGN REQUIREMENTS ...... 9

CIRCULAR BODY AIRCRAFT WITH NON-INTEGRAL TANKS (CONCEPT 1) SYNTHESIS ...... 11 Material Selection ...... 11 Tradestudies ...... 14 Structural Analysis ...... 21 CoolingSystem Analysis ...... 27

INTEGRAL TANKAGETHERMAL PROTECTION/ACTIVE COOLING SYSTEM SELECT ION ...... 33

CIRCULAR BODY AIRCRAFT WITHINTEGRAL TANKS (CONCEPT 2) SYNTHESIS ...... 37 TradeStudies ...... 37 Structural Analysis ...... 40 CoolingSystem Analysis ...... 43

7 ELLIPTICAL (BLENDED) BODY AIRCRAFT WITH INTF. GRAL TANKAGE (CONCEPT 3) SYNTHESIS ...... 45 TradeStudies ...... 45 Structural Analysis ...... 47 CoolingSystem Analysis ...... 51

8 FINALAIRCRAFT CHARACTERISTICS ...... 53

9 OBSERVATIONS AND CONCLUSIONS ...... 57

10 CRITICAL AREAS REQUIRINGADDITIONAL RESEARCH AND TECHNOLOGY . 65 Vapor Barriers ...... 65 Materials ...... 65 RadiativeThermal Protection Systems ...... 66 Subcooled LH2 Fuel ...... 66 Cooling SystemOptimization ...... 66 11 REFERENCES ...... 69

V LIST OFFIGURES

Number Page

1 Fuselage/TankStructure Study Aircraft Concepts ...... 1

2 Designand Performance Characteristics ...... 2 3 StudyPlan ...... 7

4 MissionTrajectory ...... 10

5 Summary Material Evaluation .Elevated Temperature ..... 12

6 Summary Material Evaluation .Cryogenic Temperature ..... 13

7 Concept 1 BaselineAircraft ...... 14 8 Range Sensitivity.Concepts 1 and 2 ...... 15

9 ForwardPassenger Compartment Selected ...... 16

10 Results .Fuel TankLength Study: (Elliptical Domes) .... 17

11 Results .Fuel Tank Dome ShapeStudy: (2 Tanks) ...... 17

12 ComparisonBetween Honeycomb andStiffened Structural Concepts 18

13 FlightTrajectory ...... 19

14 Effect ofAscent Trajectory on Heating Rates ...... 20

15 Non-IntegralTankage TPS Characteristics ...... 22

16 Range Penaity Associated with Changes in Non-Integral Tankage Insulation Thickness ...... 22

17 Concept 1 Structure AssemblyBreakdown ...... 23

18 Concept 1 and 2 Net AircraftShear and Moment. 2.5g Flight Condition ...... 24

19 Concept 1 and 2 Net AircraftShear and Moment. 2gTaxi Condition ...... 24

20 Concept 1 Finite ElementComputer Model ...... 26

21 AirframeSurface Active CoolingSystem ...... 28

22 TypicalCoolant Distribution Routing at Major Component Location ...... 29

...... " ...... " -.. ... LIST OF FIGURES (Continued)

Number Page

23 Simplified Distribution System Schematic. Concepts 1 and 2 . 29 24 Integration of Subsystems with Active Cooling System .... 30 25 ThermodynamicSummary. Concept 1 ...... 31

26 Candidate Integral Tankage TPS Concepts ...... 34 27 Integral Tank TPS Selection ...... 35

28 Concept 2 Baseline Aircraft ...... 37

29 Tank Wall/Primary Structure Construction ...... 38 30 Semi-structural Fuselage Covering ...... 40

31 Concept 2 Structural Assembly Breakdown ...... 41

32 Concept 2 .Finite Element Computer Model ...... 42

33 ThermodynamicSummary. Concept 2 ...... 44 34 Concept 3 Baseline Aircraft ...... 45

35 RangeSensitivity. Concept 3 ...... 46

36 TankCross Section Selection ...... 47

37 Concept 3 Structural Assembly Breakdown ...... 48

38 Concept 3 Net Aircraft Shear and Moment.2.5g Flight Condition ...... 49

39 Concept 3 Net Aircraft Shear and Moment. 2g Taxi Condition . 49

40 Concept 3 Finite Element Computer Model ...... 50

41 ThermodynamicSummary. Concept 3 ...... 51

42 FinalAircraft Characteristics ...... 53

43 FuselageITank .Group Weight Statements ...... 54

44 RelativeCost Ratios ...... 54

45 Relative Serviceability Factors ...... 55

vii LIST OF FIGURES (Continued)

Page

Structural Arrangements, S.I. Units ...... 58

Structural Arrangements, Customary Units ...... 58

Best Performance with Blended Body ...... 61

... VIII LIST OFABBREVIATIONS AND SYMBOLS

Symbol Definition Btu British Thermal Unit CG Centerof Gravity D Diameter, DragForce DW Design Weight da/ dn CrackGrowth Rate

EC Young'sModulus, Compression OF DegreesFahrenheit YieldCompressive Strength FCY FTU Ultimate Tensile Strength FTY Yield Tensile Strength ft Feet Grams, Accelerationdue to gravity GaseousHydrogen Hydrogen Hour Inch SpecificImpulse Kelvin Critical StressIntensity Factors Stress IntensityFactor L Length,Lift Force lbf PoundsForce lbm Pounds Mass

LH2 LiquidHydrogen rn Meter m Coolantflowrate M Mach number N Newton,Running Load n Load Factor N2 Nitrogen LIST OFABBREVIATIONS AND SYMBOLS (Continued)

Symbol Definition

NM Nautical Mile O.W.E. OperatingWeight, Empty Pa Pascal

PS f Poundsforce per square foot psi Poundsforce per square inch 4 Integratedheating rate = (4dA Heating rate per unit area Theoretical wing area s, sec Second T Temperature,Thrust t Thickness - t Equivalentweight thickness TO GW TakeoffGross Weight TP S Thermal Protection System W Watt, Weight P Density Prefixes

C Centi 3 k Kilo (10 ) m Milli 6 M Mega (10 ) A Difference Subscripts CR Critical F Fuel rnax Maximum S Structural s LS Sea Level Static

X Aircraft Axis - Longitudinal Y Aircraft Axis - Lateral z Aircraft Axis - Vertical

."." ...... SECTION 1 SUMMARY

A detailed analytical study was made to investigate the effects offuse- lagecross section (circular and elliptical) and structural arrangement (integral and non-integral tanks) on the performance of an actively cooled hypersoniccruise vehicle. The vehicle was a 200 passenger,liquid hydrogen fueled Mach 6. transport designed to meet a range goal of 9.26 Mm (5000 NM). Three specific point design aircraft, illustrated in Figure 1, were developed.The aerodynamic configurations were derivedfrom the NASA J3T-4 tailless delta aircraft configuration described in Reference (1). A variety of trade studies were conducted in the area of configuration arrangement,structural design, and active coolingsystem design in order to maximize theperformance of each of thethree point design aircraft. Air- craft range and weight were used as thebases for comparison and the assumption was made that adequate liquid hydrogen was available to cool the aerody- namicsurfaces of the aircraft.

MACH 6 HYPERSONIC TRANSPORT GE5/JZ6-C TURBORAMJETS

CIRCULAR WING-BODY CONCEPT 1. NON-INTEGRAL TANKS CONCEPT 2. INTEGRAL TANKS

ELLIPTICAL BLENDEDWING-BODY CONCEPT 3. MULTIBUBBLE INTEGRAL TAN

FIGURE 1 FUSELAGE/TANK STRUCTURE STUDY AIRCRAFT CONCEPTS The resultantdesign and performance characteristics are summarized in Figure 2. Concept 3, theblended body, integral tank aircraft, weighed the least andhad thegreatest range capability (over 0.47 Mm (250 NM) more than theothers). This superior performance is a resultof the better aerodynamic characteristics and higher volumetric efficiency.

CONCEPT 1 CONCEPT 2 CONCEPT 3 CIRCULAR BLENDED BODY &Z2z2 *(360.5 FT)4 -(360.5 FT)"--/ k(328.5 FT)A 109.9 rn 109.9 rn 100.1 rn TANKAGE NON "INTEGRAL INTEGRAL INTEGRAL TANK WALL MONOCOQUE ISOG R I D ISOGRID LID 4.6 4.6 4.8 VOLUMETRIC EFFICIENCY 0.67 0.71 0.88 O.W.E. Mg (LBM) 190.2 (419,200) 190.6 (420,300) 187.2 (412,800) TOGW Mg (LBM) 299.0 (659,200) 299.5 (660,300) 296.1 (652,800) RANGE Mrn(NM) 8.69 (4,690) 8.73 (4,715) 9.20 (4,968) Note: L/D Basis - Nacelle on and Cool Wall Vol eff = Tank vol/available volume in fuselage tank section

FIGURE 2 DESIGN AND PERFORMANCE CHARACTERISTICS

The relative producibility and serviceability of eachof the three aircraft concepts when optimizedfor maximum range were assessedto provide an indication of costtrends. This analysis was not as detailed as thedesign studies,but the relative ranking of these factors is believedto be accurate. Thesefactors are givenbelow, with higher numbers indicating higher cost or greatermaintenance requirements. Concept 1 is used as thebaseline for comparison. Concept 1 Concept 2 Concept 3 Isog rid Isogrid Isogrid Tank Wall Monocoque Stiffened Stiffened Producibility 1 3.5 3 Serviceability 1 1.2 1.3

2 Designapproaches which would improve the producibility of the Concept 2 and 3 aircraft were briefly examined. All theseapproaches resulted in increasedweight and therefore shorter range. Themost attractive ofthese approaches was to use monocoque tank wall construction as in Concept 1. The effect of this approach, and the associated decrease in the range of each aircraft concept is presentedbelow.

Concept 1 ConceDt 2 w Tank Wall Monocoque MonocoqueMonocoque Producihility 1 1.6 1.0 Range Loss 0 452 km (244 NM) 117 km (63 NPI)

SECTION 2 INTRODUCTION

Liquidhydrogen (LH ) fuelhas several potentialadvantages over con- 2 ventionalfuels when appliedto hypersonic flight. A particularly attractive characteristic is its highspecific impulse. Additionally, it offers a sig- nificantheat sink capacity for engine and inlet cooling, and also a potential coolingof the airframe itself. However, a significant challenge for the aircraft designer is presented by liquidhydrogen's cryogenic storage temperature, 20.3 K (-423'F),and its extremelylow density (approximately 1/12 thatof conventional JP aircraft fuel).These factors, particularly the low density,result in a unique designproblem. The large volumes required for fuel containment dictate a requirementfor aircraft shapes which provide high volume per total air- craftsurface area. Typically,blended-body and all-body shapes are attrac- tivecandidates. These shapes, combined withstructural concepts in which fuelcontainment is integral with the basic structure generally lead to high volumetric utilization. The specific major issues in this study were to evaluate the effects of fuselagecross section (circular and elliptical) and structural arrangement (integraland non-integral tanks) on theperformance of a representative Mach 6 hydrogenfueled aircraft. The entireexternal surface of the aircraft was assumed tobe actively cooled. In-depth studies were conductedon the design of theconfiguration and the active cooling system. Detailed strength analysesincluded an evaluation of the impact of fracture mechanics and fatigue on thedesign of the cryogenic tank structures. These analyses emphasizedthe development of minimum weight, long life structural concepts. They furtherprovided a firm basis for calculating the weight of the detail elements of the fuselage/tank structure in lieu of using statistical weight estimates.Configuration studies focused on approaches that would maximize volume utilization, a requirement well recognized as a dominantissue. The study was conductedwith fuel weight and payload being fixed at 108.9 Mg (240,OOo Ibm) and21.8 Mg (48,000lbm), respectively. Aircraft size, weight,and range were dependentvariables. Range was selected as a viable figure of merit.

5 The studylogic, phasing and interactions are shown inFigure 3. The study was conducted in the sequence indicated, with the non-integral tank Concept 1 designbeing accomplished first, followed by analysis of the integraltank Concepts 2 and 3. The designand performance characteristics of the baseline (or preliminary)configurations were developedand used as the basis for the aircraft trade-offand design refinement process. Generalarrangement, structural design, active coolingsystem design and mission profile trade studies were accomplishedfor Concept 1 with fixed vehiclepayload acd a missionrange goal of 9.26 Mm (5,000 NM). Usingthe results ofthese trade studies, the final sizing of the aircraft was accom- plishedand the fuel required to perform the mission was determined.This was set at 108.9 Mg (240,000 lbm). ForConcepts 2 and 3 thepayload and the fuel weight (as determined from Concept 1) were heldfixed and aircraftrange was thevariable. Sufficient hydrogenfuel was assumed tocool the aircraft aerodynamic surfaces. For all concepts,structural weights were determined as well as thetake-off gross weight (TOGW). Priorto proceeding with the design refinement process for Concepts 2 and 3 a trade-offof candidate integral fuel tank thermal protec- tionsystems was accomplished.For all threeaircraft concepts, refined designand performance characteristics were developed,and range sensitivity tovarious parameters was determined.Finally, the characteristics of the threeconcepts were comparedand evaluated and conclusions were drawn.

6 Task 2 [ StructuralAnalysis 1 DeslgnCrlterla 0 LoadslTemperature 7 0 Materlalr -l Task 4 0 Srructural r \ Concepts CONCEPT 1 Strength Aorcraft Design Analysts 1 Cnular Wlng Body Non-Integral 0 NonmtegralTanks Task 3 I Tanks I I

I a Thermal Mapplng System S1mg Task 1 Tublng Manifolds Pumps Baseline Aircraft Concepts Heat Exchanger J

Range Sensitivity 0 Configurations Comparison and Analysis CONCEPTS General Evaluation No. 1. 2 and 3 Arrangements Aerodynamlc Performance 0 StructuralWeight 0 Propulslon Task 6 0 CoolingWeight CONCEPTS 1.2. and 3 Performance a Volume Utilization 0 Misslon 0 L/D. TIW. WIS Proflle ~ L J \ I\ I Task 8 Task 9 I a Marerlals f 5 a StrengthAnalyslr CONCEPT 2 I CONCEPT3 Aircraft Design I Aircraft Design integral Tanks 0 EllipticalBlended \ / Clrcular Wmg Body Wmg Body IntegralTanks 0 IntegralTanks Task5 Task 7 f \ f \ 0 FixedPayload Active Cooling Active Cooling 0 FixedFuel Weight Syatem Integral System Analysis 0 VariableRange Tanks L J Thermal Mappmg

a System Smng I, ThermolStructural Tubing Concepts Mamfolds 0 Trade-offs Pumps \ HeatExchanger )

FIGURE 3 STUDY PLAN ,.. ., ,,_...... _...- ..- .. . SECTION 3 PERFORMANCE AND DESIGN REQUIREMENTS

Performanceanalyses for the three study aircraft concepts were accom- plishedusing the flight profile, given in Figure 4, forthe nominal design mission range of 9.26 Mm (5000 NM) . The ascent flight profile for each aircraft was constrainedby sonic boom overpressure,dynamic pressure, inlet duct pressure and heating rate limits. Ascentalong these limit lines was accomplished toreach the best (.L/D) (.Isp) condition at start of Mach 6 cruise. From this pointcruise was continued at the best (L/D) (Isp) until start ofdescent. Descent was then accomplished at Max L/D. Sufficient fuel reserves to allow loiter for 20 minutes at M = 0.8 and 12.2 km (40,000 ft)altitude were provided.Additional sea level reserve is sufficientfor one go around(5 minutes at M = 0.4). Themission performance requirements common to all threeconcepts, are: 1. Cruise Mach number = 6. 2. Payload = 21.8 Mg (48,000 lbm) . 3. Fuelweight = 108.9 Mg (240,000 lbm) (determinedfrom Concept 1 designsynthesis) including allowance for boiloff during a preflightground holdof one hour. The propulsionsystems were sized to meet theperformance requirements foreach aircraft concept. The"rubberized" engines were derivedfrom the hydrogenburning GE5-JZG wraparound stoichiometric turboramjet. A new MCAIR approachto twodimensional, horizontal ramp, external compression air indus- tion syster,zs was usedfor all aircraft. Designrequirements, guidelines and assumptions common to all aircraft were : 1. All externalsurfaces cooled to a maximum structuraltemperature of 394K C250°F) except the nacelle. An unlimitedfuel heat sink was assumed availableto absorb the entire heat load. 2. DesignLoad Factors: (a)Flight n = 2.5, -1.0 (limit) z (b) TaxinZ = 2.0 (limit)

(c)Emergency landing n = 4.5, -2.0, nx = 9.0, n = +1.5 (Ultimate) z Y-

9 3. Tank pressurization = 138 kPa (20 psi)gage limit. 4. Service Life = 10,000 hrs. 5. Fuel was LH2 contained at its normalboiling point of 20.3 K (-423’F). 6. Spacebetween structure and fuel tanks was purgedto 3.45 kPa (0.5 Psi)gage. 7. (V01ume)~/~/(PlanfonnArea) was tobe approximately the same as the HT-4. 8. Structural material was to be primarily aluminum. 9. Airframesurface heating rates were based on sustainedflight conditions. No allowance was made for flightmaneuvers. 10. Designswere based on tank pressure stabilizing the structure.

120 -

100 -

+- 80 - *- 0 z0 ’ 60 - al 73 3 ._*” .w 40 -

20 -

Mach No. FIGURE 4 MISSION TRAJECTORY

10 SECTION 4 CIRCULAR BODY AIRCRAFT WITH NON-INTEGMLTANKS (CONCEPT 1) SYNTHESIS

The finalperformance characteristics for Concept 1 were previously summarized inFigure 2.These characteristics were determined as a resultof a series of trade studies and detailed structural and cooling systems analyses ofthe aircraft. Thefollowing sections summarize the refinement studies.

Material Selection A study was conductedto select the materials tobe used for primary and secondarystructure. Primary structure was limitedto the selection of an aluminum alloy. The material selection was basedprimarily onConcept 1 requirements.The selection was maintainedfor all concepts so that material propertieswould not be a variable in the final comparison andevaluation. Thatassumption was reviewed as thedesign of Concepts 2 and 3 progressed. No specialdesign requirements were foundwhich would make the original selectioninvalid. Two considerations were paramount in selection of the aluminum alloys tobe used. Those were the 394K (250°F) maximum temperatureto which the moldlinestructure was tobe cooled and the 20.3K(-423°F) cryogenic temperature at which thefuel tanks would operate. An additionalconsideration was thedecision to assemble the tanks by kelding. This provided the best vapor seal againstleakage of the hydrogen fuel and minimized the weight of the jointsrequired in assembling the large tanks. A comparison was made of material properties for several aluminum alloys showingpromise for primary structure. A summary of thatcomparison can be found inFigure 5. Based on thiscomparison, 2014-T6 and ?075-T6 alloys were eliminatedbecause of theirsusceptibility to corrosion and stress corrosion cracking.The 7475-T761 material was notcompetitive from an elevated tem- peraturestrength standpoint and had the additional disadvantage of being availablefrom only a singlesource with resulting high cost. Low strength properties at both room andelevated temperatures eliminated the 6061-T6 material andthe T6 temperof the 2219 alloy.

11 kT= 300K (80'Fl-T = 394K (25OoF) (1O;OOO Hrs)l

Material Disadvantages

I ~~~ Susceptible to corrosion, 1 2014-T6 0.671.00 NA 1 0.92 0.84 0.990.99 1.oo exfoliation, and stress 1 corrosion cracking Low fracture toughness Good corrosion resistance, at room temperature. No 2024-T81 0.47 1.00 0.34 1 1.00 1.oo 1.001.00 1.oo goodelevated temperature elevated temperature mechanical properties fracture toughness data Low initial strength, i.e., Stable for long time I at low temperatures. No 221 9-T6 0.64 0.92 NA 0.75 0.58 0.980.63 9.86 exposure at elevated I Elevated temperature KC temperature data High fracture toughness, -t stable for long time exposurf Low initial strength, i.e., to elevated temperature. at low temperatures. 2219-T87 1.00 0.98 1.00 0.79 0.77 0.95 No 1 good corrosion resistance elevated temperature weldable, property data o.82 I 0.98 KC data readily available at elevated temperature I High fracture toughness. Low strength. No elevated 6061-T6 1 0.69 1 1.00 0.41 1 0.59 0.57 0.61 1.00 0.85 Excellent corrosion temperature KC data resistance I SusceDtible to corrosion. ' Exfoliation, andstress 1 corrosioncracking. Low 70757T6 0.91 0.76 0.84 0.82 0.80 0.98 0.95 fracture toughness. Tem- ~ 0.60 perature limited. No I ~ elevatedtemperature KC data. I Sole source, premium price, temperature limited. No 7475-T761 1.00 0.59 0.76 0.97 0.85 elevated temperature KC 0.78 0.89 data 1 I Note: Index rating ratio property to highest value in column. Highest number is best rating h.%indicates data not available FIGURE 5 SUMMARY MATERIAL EVALUATION - ELEVATED TEMPERATURE Only2024-T81 and 2219-T87 aluminum were finallyconsidered as the basicconstruction material forthe actively cooled panels. The failure modes consideredto be most significant were: (1) stability,with (Ec) com- pressivemodulus of elasticity being the figure of merit, and(2) fracture mechanics,with crack growth rate (da/dn)and fracture toughness coefficient (Kc) beingthe figures of merit. The2219-T87 alloy was shown tobe competitive in stability parameters and to have a definite superiority over 2024-T81 infracture mechanics. 2219-T87 was thereforeselected as theconstruction material forthose structural elements operating at elevatedtemperature. The choiceof materials forcryogenic tank construction was limitedby theaforementioned decision to utilize all-welded construction. A comparison ofthe three candidate materials is presentedin Figure 6. From this comparison it became obviousthat the crack growth characteristics (da/dn) of6061-T6 wouldeliminate it fromfurther consideration. The2014-T6 material was also eliminatedbecause of its inherentsusceptibility to corrosion and stress corrosioncracking. Therefore, 2219-T87aluminum alloy was chosen as the most acceptable material fortank fabrication.

Disadvantages

Susceptible to corrosion, 20 14-T6 2xfoliation. and stress corrosion cracking

~ ~~~~ High fracture toughness, stable for long time exposure to elevated temperature. Low room temperature 221 9-T87 Good corrosion resistance, strength weldable, property data readily available at elevated temperature High fracture toughness. Low strength. No 606 1-T6 Excellent corrosion cryogenic temperature resistance Kc data

Note: Index rating-ratio of property to highest value in column. Highest number is best rating

NA indicates data not available FIGURE 6 SUMMARY MATERIAL EVALUATION -CRYOGENIC TEMPERATURE

13 Annealedtitanium alloy 6A1-4V was used in some limited secondary struc- tureapplications, such as fuselagelinks and fittings, where concentrated loadsoccur and lighter structure would result. The 6A1-4V alloy was selected becauseof its favorablecombination of tensile strength, high fatigue allowables,and good fracturetoughness, compared to several commonly usedtitanium alloys.

TradeStudies The baseline(preliminary) Concept 1 aircraft characteristics are given inFigure 7. Thisbaseline met all ofthe program performance requirements and was based on estimatedweights of the fuselage/tank structure and the active coolingsystem. The fuelweight to perform the mission was determined as 108.9 Mg (240,000 lbm) which was thenfixed for the Concept 2 and 3 studies. Tradestudies were conductedto determine the effect of considerations such as passengercompartment location, tank size, construction concepts, and thermalprotection/active cooling system designs. Detail discussionsof these tradestudies are found inReferences (21, (3) and(4). In order to evaluate

PRELIMINARY CHARACTERISTICS O.W.E. = 190 Mg (419,234 LBM)

WTFUEL = 108.9 Mg (240,000 LBM) T NON-INTEGRAL FUEL TANKS 38.0 rn ( 124.8 FT) PAY LOAD = 21.8 Mg (48,000 LBM) 1 (200 PASSENGERS) ENGINE (4) GE5/JZ6-C TSLS = 400 kN (90,000 LBF) UNINSTALLED PER ENGINE

HOT NACELLE.STRUCTURE

SPEED M = 6.0(MAXIMUM)

SURFACE ACTIVELY COOLED TO 394 K (250OF) MAX

RANGE GOAL = 9.26 Mrn (5000 NM)

CONCEPT 1 BASELINE AIRCRAFT

14 trade study results in terms ofaircraft range, the sensitivity curve shown in Figure 8 was developedfor the aircraft. The resultsof these studies for the Concept 1 aircraft are presentedbelow. a. Payload/FuelLocation - Thepurpose of this tradestudy was to evaluate the effect onrange of four different passenger compartment locations: a) at a forwardposition; b) at theaircraft center of gravity (CG) withan upper tier of seats; c) at the CG with a lower tier; andd) at the CG with a shortcompartment arranged vertically in several tiers of seats. The results are presentedin Figure 9. The significanteffect of the passenger compartment location onusable fuel volume is evident. The increasedrange resulting from the better volumetric efficiency is theprimary reason for selec- tingthe forward location as the mostpromising of thefour locations examined. Inaddition, the forward location results in an ideal ground handlingarrangement for boarding and deboarding passengers and for aircraft servicing. The distinctphysical separation of thepressurized passenger compartmentfrom the fuel tank compartment also provides the best fabrication schemeof thefour arrangements examined.

270

260

E 9 250 rl E!

I .-6 240 -z J U 230

220

95 I I I I '175 180 185 190 195 190 185 180 '175 200 O.W.E. - Mg I I I I I I 390 400 410 420 430 440 430 420 410 400 390 O.W.E. - IO3 Ibm

FIGURE 8 RANGE SENSITIVITY, CONCEPTS 1 AND 2

15 --==&J=/ ..._...... :...... _...... ,._...... ,.,.,. ~ :... ;.:. ,: -A&...... :...... ,.

C /7 D "4...... :...... ,..._...... --&& .$?aI[: BASELINE: CONCEPT NO. 1 AIRCRAFT

A 0.478 0 B 0.444 -0.74 (-400) C 0.381 -2.04 (-1100) D 0.388 -1.85 (-1000)

FIGURE 9 FORWARD PASSENGER COMPARTMENT SELECTED

b.Tank Length and Dome Shape - The objectiveof this study was to determinethe combination of number of tanks and dome shapethat maximized aircraftrange. Combinations of two, three and four tanks with elliptical, torisphericaland hemispherical domes were evaluated. The results are sum- marizedin Figures 10 and 11. On thebasis of this comparison a two tank configuration with elliptically domed tankends was selected for the Con- cept 1 aircraft. c. ActivelyCooled Fuselage Covering - Activelycooled panels of both honeycomband beadedconstruction were compared forthe fuselage covering. The critical failure mode forthese panels was compressivebuckling. Coolant tubespacing was selected to limit skintemperature gradients to 56K (100°F) maximum suchthat thermal stresses had a negligible effect on panel general stability.Weights were determinedfor the panels considering not only the loadcarrying structure but also splices, manifolds, shear clips, adhesives, fasteners,residual coolant and a pumpingpower penalty. The studyresults are summarized inFigure 12. Thehoneycomb panelconstruction is lighter, withinthe Concept 1 had range, andprovides for a greaterrange potential.

16 FIGURE 10 RESULTS-FUEL TANK LENGTH STUDY: (ELLIPTICAL DOMES)

FIGURE 11 RESULTS-FUEL TANK DOMESHAPE STUDY: (2 TANKS)

Theseadvantages, combined with the other desirable features listed in Figure 12, ledto the selection of honeycomb panelconstruction for the fuselagecovering.

17 Actively Cooled Actively Cooled Honeycomb Corrugation Stiffened Fuselage Primary Structure Panel-Stringer Concept

Fuselage Frame Fuselage 1 Insulation -, Actively Cooled Frame Honevcomb 7

Tank Wall

Panel Unit Weight vs Ultimate Allowable Load - Selected Honeycomb Panels 0 74.1 km (40 NM) RangeSaving 16 (v RangeLoading 0 Lighter E - ". Containment Fluid Fail-safe 3 14- Minimize Fastener Installation

4- Utilization Structural0 More Efficient L 07 with Frame Caps - g12- .-i- c 3 - 10 I I 1 I 100 200 300500 400 NCR - kN/m I I I 500 1000 1500 2000 2500200015001000 500 3000 NCR - Ibf/in.

FIGURE 12 COMPARISON BETWEEN HONEYCOMB AND STIFFENED STRUCTURALCONCEPTS

d.Trajectory Shaping - A studyrelated to the ascent trajectory provided visibility to the sensitivity of this class of aircraft to the maximum heating rate encounteredin the flight profile. As indicatedin Figure 13, theorigi- nally assumedascent was based solely on sonic boom overpressure,dynamic pressure,and inlet duct pressure limitations. The design of theactive coolingsystem to this profile required the absorption of a transientheat load

18 120 - 35 - 0 Pressures are absolute values. Mach 6 Cruise 100 - 30 - Constant Heating

0 NumberClimb 80 - 25 - c.' .I- O 0 E z Y 20- Constant Dynamic Pressure 0 '60- m =71.8 kPa (1500 Ibm/ft2) 0) U 896 kPa (130psi) U 3 3 Inlet Duct Pressure .-4d 'E 15- 4d 3 ' 40- 10 - - -- Original Trajectory -Final Trajectory 20 - 5-

0- 0 1 2 3 4 5 6 Mach Number

FIGURE 13 FLIGHT TRAJECTORY

that was 40% greaterthan the sustained heat load experienced during cruise as indicatedin Figure 14. This initial approach resulted in a large,heavy activecooling system which was significantly overdesigned for mostof the mission,including all ofcruise. A modificationto the ascent trajectory, following a constantheating pathfrom Mach 5 to the Mach 6 start ofcruise at thebest (L/D)(Isp) condition, was theninvestigated. This approach resulted in a netrange gain of 289 km (156 NM) attributableto a reductionin cooling system weight. This ascent trajectory modification was incorporated in the final trajectory. e. Nacelle Cooling - The feasibility ofcooling the nacelle module surfacesto the same temperature, 394 K (250°F), as the rest ofthe airframe was studied. It was foundthat although the surface area involvedrepresented only9.4% of thetotal airframe wetted surface area, coolingthe nacelle sur- facesadded 23.8% to the total heat load to be absorbed.

I 1.5 - Mach x,[ OriginalTrajectory / Mach 5 Mach 6 1.0 -

Final Trajectory

Mach 4

0.5 -

0 I r I I 0 20 25 30 35 Altitude - km ! 0 60 12080 100 Altitude - 1000 ft

FIGURE 14 EFFECT OF ASCENT TRAJECTORY ON HEATING RATES

There are numerousreasons why theheating rates onthe nacelle surface were much higher than the overall average surface heating rates. o The nacelle is located on thelower surface.

o Most nacellesurface locations are nearboundary layer origins which results in high heat transfer rates due to the short characteristic lengthsinvolved. o Flow inthe boundary layer diverter region is subsonicand therefore localadiabatic wall temperaturesapproach total temperature. o The externalinlet ramps are positioned at highdeflection angles. o Heatingto the panels in some regionsincludes conduction from the internal duct walls in addition to external aerodynamic heating. All of thesefactors combine to imposeextreme cooling requirements for the nacellesurfaces. A comparisonwith a "hotf'structure nacelle design was made. Superalloy materialscompatible with the resultant higher temperature environment were used. Whilethe nacelle structural weight increased, a netaircraft system weightdecrease of 2.39 Mg (5266lbm) resulted,corresponding to a 137 km

20 (74 NM) gainin range. In addition, cooling the nacelle surface results in many practicaldesign problems. Routing coolant lines across the fuselage/nacelle interfaceand designing cooling lines into the external inlet duct rampsand sidewallswould be complex and probably would result in volumetric penalties. Taking all ofthese factors into account, the "hot'' structure approach was selected. f. HydrogenTankage Thermal Protection - Thermalprotection for the Concept 1 non-integraltankage consists of cooled surface panels, a nitrogen purgedgap, and foam insulation applied to the external tank surfaces. A study was conductedto find the combination of insulation weight and fuel boiloffweight that maximized aircraft range. This study was based on the missionrequirements which include a onehour ground hold preceding the flight as well as a 20 minuteloiter at 12.2 km (40,000ft) prior to descent. Parametricdata on weighteffects alone indicate that the minimum combined weightof insulation and fuel boiloff occurs with an insulation thickness near 2.54 cm (1.0 inch)as shown inFigure 15. However, it was determinedthat achieving more usablefuel at theexpense of additional insulation weight was beneficialfrom a rangestandpoint. Maximum range was determinedto be achievedwith an insulation thickness of 4.27 cm (1.68in.) as shown in Figure 16.

Structural Analysis The structuralarrangement for Concept 1 is shown inFigure 17. Inorder toinitially determine the vehicle size and performance capability, an "initial estimatedweight" was determined.This weight was based on current MCAIR esti- mationtechniques modified for use with actively cooled structure. Using these techniques,factors were appliedto the forward fuselage pressurized passenger compartment toaccount for the non-circularity of the fuselage cross section shape.Weights for all structural componentsexcept those located in the fuselageltank(center fuselage) region of the aircraft, as determinedwith

21 6- 1.2 -

5- 1.0 - Insulation + Boiloff

0.8 - 4- Insulation N N .r E E .m B Y a 0.6- ;3- L s m .-03 ._ 2 3

0.4 - 2- -Fuel Bolloff. Flight

0.2 - 1-

0- 0 I I I I 0 2 4 6 8 Insulation Thickness. crn

I I I I 0 1 2 3 Insulation Thickness - in.

FIGURE 15 NON-INTEGRAL TANKAGE IPS CHARACTERISTICS

2o 1 10 -

8- r E z Y 6- c .- ._ .-u r u- d 0" W W m 4- c m a d

I1 68 in.) 2- \=4';cm I 0- 0 2 4 6 Insulation Thickness. cm

I I I I I I 0 0.5 1.0 1.5 2.0 2.5 Insulation Thickness - in.

FIGURE 16 RANGE PENALTY ASSOCIATED WITH CHANGES IN NON-INTEGRAL TANKAGE INSULATION THICKNESS

22 Note: Active cooled panels cover all external surface except nacelles.

Fwd Tank

,/'

Upper Center Fuselage

IU IUA Ill IU .A - Fuselage

Lower Center Fuselage

I Aft Wing

\-PassengerCabin Fwd Wing

FlGlJRE 17 LNacelle Module CONCEPT 1 STRUCTURE ASSEMBLY BREAKDOWN thesetechniques, then remained fixed. Detailed structural analyses of representativestructural components located in thefuselage/tank region were thenconducted to refine the center fuselage structure and fuel tank weights. The resulting refined weights were then used to determine the final aircraft rangeand permit performance comparisons with the other aircraft concepts. Theweight refinement process started with the definition of a finite elementcomputer model of the fuselage tank area which was utilized to determine internal load distributions resulting from overall aircraft loads. Althoughonly the fuselage tank area was modelled,stiffness and load characteristics ofadjacent structural regions were consideredin the model. Because ofthe inherent stiffness characteristics of the monocoque constructionof the forwardand aftfuselage sections, it was logically assumed thatload introduc- tionfrom these elements would take the classic plane strain distribution. Inputload vectors were thereforecalculated on thatbasis. Wing loadinputs tookthe form of cap loads and web shearsin each of the spars. A simplified quickanalysis indicated that the chordwise bending stiffness of thewing did notcontribute significantly to the overall fuselage bending stiffness and was thereforeneglected. Design loads used in the Concept 1 analysis are presentedin Figures 18 and 19. The internalloads were thenemployed instrength analysisto determine member sizingand to assess theeffects of fatigue and fracturemechanics requirements. This sizing analysis process was automated usingthe MCAIR ComputerAided StructuralDesign (CASD) program. a. FiniteElement Computer Model - The structural modelused foranalysis of theConcept 1 centerfuselage is illustratedin Figure 20. With thismodel, 1055joint degrees of freedom were used. A resizingroutine was usedwith threeiterations to obtain margins of safetynear the desired zero value. b.Fuselage Structure - Detailed stress analysis of theactively cooled panels,frames, bulkheads and longerons which form the fuselageltank area primary structure resulted in a calculatedweight of16.41 Mg (36,185 lbm') comparedwith the initial estimated weight of 18.16 Mg (40,037 lbm). In the area ofthe monocoque structuralshell, the actively cooled panels were heavierthan originally estimated. However, theframe and bulkhead weights were lighterthan estimated, largely as a resultof the use of the panels as partof the frame caps. Accomodatingthe N purgepressure requirement 2

24 Mach 3.0 Altitude = 15.2 km (50,000ft) rlZ = 2.59 DW = 299.02 Mg (659,234 Ibm) Concept 1 DW = 299.48 Mg (660,252 Ibm) Concept 2 1.01 0.8 - - 10

-0.2 - -0.4 - -0.6 - - -8 -0.8 - "- -1 0 -1 .o I I I I I I I I I -12 0 20 10 30 40 50 60 70 80110 100 90 Fuselage Station . meters I I I I I 1 I I I 0 40 80 120 320 280160 240 200 Fuselage Station. ft

FIGURE 18 CONCEPT 1 AND 2 Net Aircraft Shear and Moment 2.59 Flight Condition

3.0 6- 2.5 5- 2.0 4- -8 1.5 3- -6 1.0 - -4

1 - 0.5 - 2, r E D D 0--, 0 - o.= ro I z (0 -1 - -0.5 s - -2 0

L m c - -4 5 .c2 -2 - 2 -1.0 v) $ -3 - - -6 I -1.5 4- - -8 -2.0 -5 - - -10 -2.5 -6 - - -12 -3.0 -7 - - -14

-3.5 "-16

-4.0 1 I I I I I I I I I I I -24 0 40 10 30 20 50 60 70 80 90 100 110 Fuselage Station . meters I I I I I I I I I 0 40 80160 120 200280 240 320 Fuselage Station . ft

FIGURE 19 CONCEPT 1 AND 2 Net Aircraft Shear and Moment 29 Taxi Condition

25 PAN E LS - 423 0 1055 JOINT DEGREES OF FREEDOM

FIGURE 20 CONCEPT 1 FINITE ELEMENT COMPUTER MODEL Fuselage/Tank Area

resultedin only a modestweight effect, specifically 44 kg (97 lbrn), which affectsrange by only 2.5.3 km (1.4 NM). The effects of thefatigue and fracture mechanics requirements were also found to be small as noted:

o Weightincluded for fatigue = 0

o Weight includedfor fracture mechanics = 268kg (590 lbm), equivalent to14.8 km (9 NM) range. c. FuelTanks - A plainskin, monocoque construction was selectedfor use on theConcept 1 non-integraltanks. The skinthickness was established by burstpressure requirements. The pressure stabilized tanks showedadequate margins of safety€or the crash and flight bending conditions. Use of stiffening on thetank walls was notrequired once the monocoque thickness was establishedby pressure requirements. Theunique combination of tank dimen- sionsand pressure were completelyresponsible for the selection of monocoque tank walls from a weightstandpoint. Analysis of the two non-integraltanks ledto a calculatedweight of 7.12 Mg (15,699lbm) as compared to the initial estimatedweight of 7.09 Mg (15,635lbm).

26 Cooling- "" System Analysis A detailedstudy of the aircraft's active coolingsystem was conductedto refinethe system weight and provide a completesystem definition for comparisonwith the other study aircraft. As indicated in Figure21, the active coolingsystem uses coolant passing through the structural surface panels to absorbthe aerodynamic heating input to the airframe. A distributionsystem routes coolant to and from the panels , passingthrough a heatexchanger where theheat load is transferredto the hydrogen fuel. The panels are heldto a maximum surfacetemperature of 394 K (250°F). The coolant inlet temperature tothe panels was assumed to be 256 K (0°F).Coolant tube spacing was chosen topermit a 56 K (100°F) outerskin thermal gradient and minimize panel structuralweight. As a result,the average coolant outlet temperature was approximately 292 K (65°F).Sufficient hydrogen fuel heat sink capacity for airframe cooling was assumed,as part of theinitial study ground rules. It wasassumed thatthe LH2 fromthe tanks was suppliedto the system heat exchanger at 20.3 K (-423°F).Subsequent analyses, based on a fueltemperature rise of 235 K (423°F) , revealed that engine fuel f lowrates during cruise would, in fact, not providesufficient heat capacity to absorb the total air.frame heatload. These flowratesare only slightly more thanhalf of thatrequired. This subject is discussedrelative to each concept in subsequent sections of this report. Figure 22 indicates how coolantfeeder lines, branching out of the main lines,service adjacent surface panels. Coolant is dispersedinto manifolds atthe panel ends to distribute flow evenly through coolant tubes. A distributionsystem routing, indicated in Figure 23, was derivedbased on 6.1 m (20 ft)long panels with the fuselage panels positioned in a staggered arrangement,which reduces' feeder line size. Aircraftsubsystems (environmental control, hydraulic, and electrical systems) were definedonly in sufficient detail to establish their cooling requirements.These systems were integratedwith the airframe's coolant distributionsystem, as shown inFigure 24, so thatthese heat loads are alsoabsorbed by the hydrogen fuel.

27 Typical Actively Cooled Surface Panel

tTmax = 394 K (250°

Circumferential Return Feeder Aerodynamic Heating

Tin. = 256 K (OOF) Main Coolant Return Line Note: Temperatures are Representative of Start of Cruise at Mach 6 FIGURE 21 AIRFRAME SURFACE ACTIVE COOLING SYSTEM

28 PANEL JOINT

CONNECTION AT PANEL MAN1FOLD - TYPICAL

CIRCUMFERENTIAL RETURN FEEDER LINE

CIRCUMFERENTIAL SUPPLY FEEDER LiNE

H2 FROM FUEL TANK

MAIN SUPPLY LINE - R/H

MAIN SUPPLY EXCHANGERLINEHEAT - L/H FIGURE 22 TYPICAL COOLANT DISTRIBUTION ROUTING AT MAJOR COMPONENT LOCATION UPPER CIRCUMFERENTIAL FEEDER LINE

S IDE CIRCUMFERENTIALSIDE VERTICAL TAIL F EE D ER LINE DISTRIBUTIONLINE LINE FEEDER

MAIN SUPPLY LINE

LOWER CIRCUMFERENTIAL FEEDER LINE HEAT EXCHANGER

AFT FUSELAGE MAIN SUPPLY LINE

Supply Lines Only are Shown, Return Lines Spaced Similarly Lines Represent Those on One Side of Aircraft Only, Exception Vertical Tail Lines FIGURE 23 SIMPLIFIED DISTRIBUTIONSYSTEM SCHEMATIC, CONCEPTS 1 AND 2

29 .

Aft Fuselage Main Supply Line Main Supply LineForwardSupply Main Fuselage LineSupply Main .

WaterCoolant I - H2 from +""" Tanks Air to Pressurized Compartment Cooling System Heat Exchanger

H2 to t" Engines Air From Pressurized Pumo Compartment Package I Pump A Leakage Air Make-up uHydraulic System Intermediate Electrical Coolant System Exchanger System Exchanger

Intermediate Coolant J I Aft Fuselage ' Forward FuselageMain Return Line &5 Main Return Line -Coolant - - - - HydrogenFuel "- Air 0ValveControl

FIGURE 24 INTEGRATION OF SUBSYSTEMS WITH ACTIVE COOLING SYSTEM

Surfaceheating rates were established at thecooling system design point selectedfrom the ascent trajectory trade study. These heating rates varied from maximuma of 199 kW/m2 (17.5 Btu/secft2) on theforward fuselage to a minimum of 12 kW/m2 (1.1Btu/sec f t2) on theaft fuselage upper surfaces. The averageheating rate overthe entire cooled surface area was 29.4 kW/m2 (2.59Btu/sec ft2). Parametric data generated with a computerizedthermal model of a cooledpanel were usedto establish flowrate requirements. This information, combined withthe distribution system definition, enabled the activecooling system weights to be determined. Figure 25 provides a summary of theresults of thermodynamic analyses of theConcept 1 aircraft. Heat loadsand coolant requirements for each majorsection of the aircraft are shown alongwith totals which include subsystemrequirements. A coolingsystem weight breakdown is providedand thefuel tank thermal protection system characteristics are summarized.These results are based on thespecified assumption of unlimited fuel heat sink capa- cityto cool the entire aircraft surface (exclusive of the nacelle). As

30 Hydrogen Fuel Tankage Thermal Protection System

hertical tail =r 9.3 MW (8.78 X lo3 Btu/sec) m,ertical tail= 64 kg/s (142 Ibm/sec)

Insulation Weight = 2.79 Ma (6.160Ibm) 12.7 cm (5in.) Fuel Boiloff

Environmental Control System and Purge System Components

Qfuselage = 50.4 MW (4.78x lo4 Btu/sec) Active Cooling System Weight mfuselage= 401 kg/s (883Ibm/sec) Component: Mg (Ibm) Residual Coolant (21,130) 9.58 ActiveCooling SystemHeatExchanger Distribution Lines, etc. 1.46 (3,210) ’ Exchanger Heat 1.13 (2,500) 1.13 qota1= 108.5 MW (1.029 x IO5 Btu/sec) Pumps and Pump Fuel Req(6.595) 2.99 mtotal = 848 kg/s (1,869Ibm/sec) Total 15.17 (33,435)

Note: Totals include subsystem requirements. FIGURE 25 THERMODYNAMIC SUMMARY, CONCEPT 1

mentionedpreviously, actual engine fuel flowrates during cruise are inadequateto absorb the total heat load. For Concept 1, it was estimatedthat theheat sink afforded by the fuelflow during cruise is approximately SO% of thatrequired to absorb the design heating rate. However,refinement of the airframethermal protection system via localized heat shielding, etc., to provide a matching of theairframe heating rates andengine fuel flow heat capacitiescompatible with the engine efficiencies used was consideredto be beyondthe scope of thisstudy. Themajor goal of thisstudy was establish- mentof a baselinesystem against which fuselagejtankage trade studies could be made.

SECTION 5 INTEGRAL TANKAGE THERMAL PROTECTION/ACTIVE COOLING SYSTEHSELECTION

Priorto the design refinement of Concepts 2 and 3, a study was conducted to select a thermalprotection/active cooling system arrangement for the integraltank aircraft concepts.Trade studies of eight candidate arrangements were conductedand the most promising arrangement was selected based on the fact that no light weight vapor barrier non-permeable to gaseous hydrogen exists. The eightcandidate conceptual arrangements considered are shown in Figure 26. Concepts @ , @ , and @ were specifiedin the study definition. Concepts @ through @ were evaluatedto insure that a varietyof competitive arrangements were investigated.Concept 0,due to its thermodynamic simi- larityto the non-integral tankage thermal protection system (TPS) arrangement, was selected as thebaseline concept. Range differences were determinedto reflectdifferences in TPS weight,usable fuel, and TPS volume. The range differencesreflect configuration variations on theupper half of the fuselage only.This simplification avoided numerous complexitiesinvolved with con- siderationsunique to the lower half of the fuselage as justifiedin Refer- ence (3) . Therefore,the range differences shown are approximatelyone half theactual magnitude involved. Eachconcept was alsoevaluated in terms of relative fabricationdifficulties, inspectability/maintainability, cost,and development status. A summary of thisinformation is presentedin Figure 27. As notedon the figure, Concept 0,the baseline, was alsoselected as the integraltankage TPS concept. It canbe noted in Figure 27 that most of thecandidate arrangements are reasonablycompetitive on a rangebasis or purely on a unitweight basis. Therefore,other considerations were primedrivers in the final configuration selectionprocess. Concepts @ and @, whichrequire thick layers of insulation due to H2 gaspermeation, penalize the usable fuel volume to theextent thatthe resultant range losses are significant.Concepts @ and @ require diffusion bonded structureto insure against hydrogen leakage. Such structures would be expensiveand difficult to inspect. Concept @was considereddifficult tofabricate and inspect. Concept@, althoughanalytically attractive, would requirethe development of an acceptable multilayer, evacuated insulation material. Finally,Concept @ offersno significant advantages over Con- cept @ and is less thermodynamicallyefficient. Hence, Concept 0 was selectedfor subsequent studies.

I ~ I Concept @: Internal Insulation l- Concept @: Internal and External Insulation

e Primary System Coolant Primary System Coolan e Panel Inner Skin Serves Purged Locally Around as Tank Wall Coolant Feeder Lines e Purge Locally Around Non-Permeated Insula- Coolant Feeder Lines tion e GH2 Permeated Stiffened Tank Wall Insulation e GH2 Permeated Insulation Coolant Feeder Line (Typ)

- ~~ Concept @: Hydrogen Cooled Surface Concept 0:Internal and External Insulation/ Panels/lnternal Insulation Hydrogen Boil-Off Cooled Structure

Primary System Coolan. Purge Locally Around Direct Hydrogen Cooled Coolant Feeder Lines Panel Inner Skin Serves Non-Permeated Insula- as Tank Wall tion No Purge Requirement Hz Boiloff Heat Ex- GH2 Permeated changer Inner Skin Insulation Serves as Tank Wall GH2 Permeated Insulation

~ 1 Concept 0:External Insulation/Gap I Concept a:Internal Insulation/Metallic Liner

e Primary System Coolani e Panel Inner Skin Serves as Tank Wall e Purge Locally Around Coolant Feeder Lines e Non-Permeated Insula- tion e Metallic Liner

Concept @: InternalInsulation/Gap Concept 0:External Multilayer, I I Evaculated Insulation/Gap

Primary System Coolant e PurgedGap Multilayer Evacuated Insulation Stiffened Tank Wall

FIGURE 26 CANDIDATE INTEGRAL TANKAGETPS CONCEPTS

34 112 RANGE i CHANGE FROM DIFFERENCESSELECTEDFROM CONCEPT STRUCTURE t BASELINE TPS UNIT CONCEPT @ FABRI- INSPECTABILITY/ FIXED WEIGHT CONCEPT km(NM) RANGE COST DEVELOPMENT MAINTAINABILITY , kg/m2 (lbm/CATION f t 2, STRUCTURAL COOLED PANEL, -352 (190) SIGNIFI-DIFFICUL'L' POOR I HIGH 1 ------23.7(4.86) ! 'PERMEATED INSULATION LOSS1 CANT 1 9STRUCTURAL H2 COOLED PANEL , "" POORDIFFICULT 22.0(4.50) PERMEATED INSULATION* LOS s DIFFICULT POOR 14.7(3.02) I1 1I I I 3 STRUCTURAL TANK WALL, BASELINE AND SELECTED CONCEPT 19.4(3.98) NON-PERMEATED INSULATION, GAP

3 STRUCTURAL TANK WALL, -174(94) SIGNIFI- "" "" -"- 20.9(4.28) PERMEATED CANT LOSS INSULATION, GAP

3 STRUCTURAL TANK WALL, -44 (24) LOS s "" "" "" 20.5(4.19) NON-PERMEATED INSULATION 3 STRUCTURAL BOILOFF Hz -98 (53) LOSS DIFFICULT POOR HIGH 24.0(4.91) COOLED TANK WALL, PERMEATED/NON-PERMEATED INSULATION g STRUCTURAL COOLED PANEL, -7 (4) "" DIFFICULT POOR HIGH 19.3(3.96) NON-PERMEATED INSULATION, METALLIC LINER

9STRUCTURAL TANK WALL +39 (21) INCREASE "" "" "" 19.5(4.00) MULTILAYER, EVACUATED INSULATION, GAP - insulationthickness same as for Concept@ As a result, fuelboiloff is inadequateto cool surface panels.Concept @ insulationthickness sized to provide adequate boiloff for structural cooling.

FIGURE 27 INTEGRAL TANK TPS SELECTION

SECTION 6 CIRCULAR BODY AIRCRAFT WITH INTEGRAL TANKS (CONCEPT 2) SYNTHESIS

The Concept 2 aircraft is almostidentical in external lines to Concept 1. It also has the same fuelweight, passenger payload and propulsion system. However, thestructural arrangement of the fuselage/tank area is quitedif- ferent, in that a singleintegral fuel tank replaced the non-integral tanks of Concept 1. Selectionof the tank thermal protection and active coolingsys- tems forConcept 2 was completed prior to refining the design of the airplane. The refinementstudies were similar tothose of Concept 1.

TradeStudies The baseline(preliminary) Concept 2 aircraftcharacteristics are given inFigure 28. The airplane size was established bythe payload and fuel weightrequirements. Range was a "fallout" of thoseassumptions. Many of thetrade study results discussed in Section 4 alsoapply to Concept 2. Pas- sengercompartment location, elliptically domed tankends, and a two compartment fueltank are samples of items retainedfrom Concept 1. The tanklength PRELIMINARY CHARACTERISTICS O.W.E. = 189.7 Mg (418,106LBM)

WTFUEL = 108.9 Mg (240,000LBM) INTEGRAL FUEL TANKS

" I PAYLOAD = 21.8 Mg (48.000 LBM) (200PASSENGERS) - ENGINE (4)GE5/JZ6-C TSLS = 400 kN (90.000 LBF) UNINSTALLED PER ENGINE

HOT NACELLE STRUCTURE

SPEED M = 6.0 (CRUISE)

SURFACE ACTIVELY COOLED TO 394 K (25OOF) MAX

FIGURE 28 CONCEPT 2 BASELINE AIRCRAFT

I tradestudy showed that the twocompartment tank arrangement was necessaryfor CG controland to limit the effect of crashcondition pressure heads. The honeycomb constructionactively cooled panels were also retained in order to provide minimum fuselageweight. The modifiedtrajectory was utilized for all of the aircraft concepts, as well as thehot engine nacelle moduleconfigura- tiondescribed in Section 4. The two majoradditional trade studies conducted for Concept 2 involvedtank wall constructionand the structural arrangement to beused for the fuselage covering in the tank area. Range sensitivityof theConcept 2 aircraft was foundto be the same as Concept 1. Thereforethe sensitivitycurve of Figure 8 was usedto evaluate these trade study results. a. .Tank Wall Construction - Non-stiffenedtank walls were found tobe a heavyapproach for the Concept 2 integraltanks due to the higher bendinE loads.Integrally machined stiffeners were chosenover mechanically fastened stiffenersbecause of the problem associated with leakage of gaseous hydrogen throughthe fastener holes. Thetwo most attractive externalstiffener arrangements, a 0"-90" waffle patternand Isogrid construction, are compared inFigure 29. Based on theseresults, the Lsogrid construction was selected forConcepts 2 and 3 torealize the weight savings potential.

600 I

400

5 6 7 89 Unit Weight - kg/rn2 Selected I I 1.0 1.81.2 1.6 1.4 Unit Weight - Ibrn/ft2

FIGURE 29 TANK WALL/PRIMARY STRUCTURE CONSTRUCTION Externally Stiffened

38 Three othermethods of tank wall construction were alsoconsidered for this study. The first was a 0-90" wafflepattern modified by the addition of discreterings. These rings would also be used as supportsfor the actively cooledpanels. Thermal gradients,however, from 20.3 K (-423OF) at thetank wall to 366 K (200'F) at thepanel inner surface would create drasticthermal stress problems in a continuousring. For this reason the rings were elimin- atedfrom further consideration. Integral stiffening by means of -+45" waffle patterns and plain monocoque skin construction were also considered but were notcompetitive from a weight standpoint. Many otherapproaches to integral stiffening are available.This study was limitedto the concepts noted above because of their simplicity, limitedheight, andready adaptability to application of cryogenic insulation. b.Semi-structural vs Non-structuralFuselage Covering - Sincethe integral tank is theprimary structural load path in the fuselage/tank area, theactively cooled panels initially were designedto be non-structural. How- ever,the panels, as designedto the minimum heightrequired to serve thecool- ingfunction, weigh approximately 90% as much as theywould in a structural configuration.Therefore, a tradestudy was conductedto determine how much weightcould be saved by using the panels as secondarybending structure. The non-structuralpanels were assumed tobe supported individually from the integraltank and have slip joints around their periphery to allow relative motion betweenthe panels. In the semi-structural arrangement all thepanels were assumed tobe interconnected andsupported, on frames, from the upper wing surface.Major slip-joints were utilizedat each end of the cover to allow thermallyinduced motion between the tank and cover. The semi-structural panels did not have to be increased in weight to carry the maximum compressive runningloads. By usingthis arrangement, the integral tank bending loads are relievedsufficiently to permit a significanttank shell weight reduction. The results showed that 998 kg (2200 lbm) intank weight could be saved, which correspondsto a rangeincrease of 51 km (31 NM) . The semi-structuralfuse- lagecover arrangement selected for Concept 2 is illustrated in Figure 30.

39 Cover Slip Joints

FIGURE 30 SEMI-STRUCTURAL FUSELAGE COVERING

Structural Analysis The structuralarrangement for Concept 2 is shown inFigure 31. The processof structural weight refinement described for Concept 1 was also followedfor Concept 2. Design loads used in the Concept 2 analysis were the same as thoseused for Concept 1 as presentedin Figures 18 and19. a. FiniteElement Computer Model - The Concept. 2 structuralmodel is illustratedin Figure 32. 1618 jointdegrees of freedom are used with this model. The fuselageand tank structures have been separated for this illus- tration,but are joinedby the computer analysis program. As forConcept 1, a resizing routine with three iterations was employed toapproach a zeromargin of safety. b.Fuselage Covering - The loadsin the semi-structural covering are approximately 20% or less, of thosefor the primary structural covering on theConcept 1 aircraft. Detail stress analysisof the semi-structural activelycooled panels resulted in a calculatedweight of 9.31 Mg (20,540lbm) comparedwith the original estimate for the baseline of 8.90 Mg (19,625lbm). No additionalweight was required to satisfy either the fatigue or fracture mechanicsrequirement.

Accommodating the N2 purgerequirement resulted in only a modestweight effect, as was the case withConcept 1. Specificallythe weight effect was anincrease of 44 kg (97 lbm) which transforms to a rangeincrement of 2.53 km (1.4 NM) c. Tank - InConcept 2 the structure was interconnectedwith link systems that accommodate thethermal strains while still maintainingreliable struc- turalload paths. As anexample the wing-to-tank connection is made with a series of linkswhich have monoball bearings at eachend to allow the links to swivel and yetallow one end of the link to move with respect to the other. Each linkhas full axial loadcapability. A series ofnearly vertical links is usedto attach each side of thefuel tank to the upper surface of the wing. The longitudinallocation of thetank is fixedby a single aft link which attachesto the wing carry-through at the centerline.Side motion of the tank is preventedby a series oftransverse links. Thermalcontraction of the tank is accommodatedby theselinks, which travelin an arc andinduce a bending stress ofonly about 3.45 MPa (500 psi) inthe tank at thepeak of the arc. Trussnetworks formed of the same typeof links provide thermal strain relief at thesplice joints where the forward and aftfuselage sections and the vertical tail attachto the tank. Testingof the Isogrid stiffening concept has shown biaxiallyloaded structure stress concentrationfactors to belimited to a maximum of 1.5. Thisvalue wasemployed inthe integral tank analysis. Detail analysisof the Concept 2 integraltank resulted in a tankweight of11.03 Mg (24,309lbm) compared with the initial estimate of8.76 Mg (19,303lbm). This analysis also resulted in a finalweight of theframe, bulkheadsand longerons of 3.61 Mg (7,955lbm) compared with the initial estimate of5.51 Mg (12,150lbm) . Althoughthese differences between the individ- ual initial and final estimates are significant,the totals were in good agree.ment.Weight added specifically for fatigue considerations was minor, 20.4kg (45 lbm) , andnone was added forfracture mechanics.

Cooling Sys tem Analysis The.Concept 2 coolingsystem was analyzed in a manner similar to that describedfor Concept 1. SinceConcept 1 and 2 moldlinecontours and areas are nearlyidentical, the surface heating rates are similar andthe cooling

43 systemrequirements are essentiallyequal. The most significantdifference betweenthe cooling system designs involved a minorrelocation of the cooling systemheat exchanger. This relocation, necessitated by thelack of available spacebetween the tank wall andthe outer skin, did not significantly impact coolingsystem weight. A summary ofthe results is presentedin Figure 33. As withConcept 1, theengine fuel flowrates during cruise are sufficient to cool about 50% of theairframe heat load.

Hydrogen Fuel Tankage Thermal Protection System

Actively Cooled Qvertic-1 tail = 9.3 MW (8.78X lo3 Btu/sec) SurfacePanel mvertical tail = 64 kg/s (142Ibm/sec)

Insulation Weight = 2.76 Mg (6,076Ibm) Fuel Boiloff = 2.56 Mg (5,642Ibm)

Q,ings = 50.7 MW (4.8x lo4 Btu/sec) m,,,,ings m,,,,ings = 401 kg/s (883Ibm/sec)

Environmental Control System and Purge System Components

&uselage = 48.1 MW (4.57 X lo4 Btu/sec) ActiveCooling SystemWeight mfuselage= kg/s Ibm/sec) Component: Ms (Ibm) - 392 (864 Residual Coolant (21,182) 9.61 Active Cooling System Heat Exchanger Distribution Lines, etc. 12)(3.21.46 .. . Heat Exchanger (2,5011.13 1 Pumps and Pump Fuel Req(6,613) 3.00 Total (33,508)15.20 Note: Totals include subsystem requirements

FIGURE 33 THERMODYNAMIC SUMMARY,CONCEPT 2

44 SECTION 7

ELLIPTICAL (BLENDED) BODY AIRCRAFT WITH INTEGRAL TANKAGE (CONCEPT 3) SYNTHESIS

The Concept 3 aircraft, with its blendedwing body configuration, while similar in planform, is quite different in cross section from the previous studyaircraft. Its ellipticalfuselage cross section resulted in a different approachto configuration of the integral fuel tank. It did,however, have the common basisof fuel weight, passenger payload, and propulsion system that were usedon Concepts 1 and 2. Tank thermalprotection and the active cooling systemsfor integral tank aircraft(Concepts 2 and 3) were discussedpreviously. Refinementof the design into the final configuration was accomplished in the same manner as previouslydescribed for the Concept 1 and 2 aircraft.

TradeStudies The baseline(preliminary) Concept 3 characteristics may befound in Figure 34. Payloadand fuel weight requirements sized the aircraft, with the aircraftrange as a "fdl out".Passenger compartment location, elliptically PRELIMINARY CHARACTERISTICS O.W.E. = 187.4 Mg (413,204 LBM)

A WTFUEL = 108.9 Mg (240,000 LBM) I INTEGRALTANKSFUEL 36.1 m (l18*4 FT) PAYLOAD = 21.8 Mg (48,000 LBM) 1 1 (200 PASSENGERS) ENGINE (4) GE5/JZ6-C TSLS = 400 kN (90,000 LBF) UNINSTALLED PER ENGINE

HOT NACELLE STRUCTURE

21.8 m SPEED M = 6.0 (MAXIMUM)

::::::x: :::::: ::..,-- - - ~ (71zFT) (.. ...- 1 SURFACE ACTIVELY COOLED TO 394 K (25OOF) MAX

FIGURE 34 CONCEPT 3 BASELINE AIRCRAFT

45 domed tankends, and a two compartment fueltank are samplesof items retained fromConcept 1. The tanklength trade study showed thatthe two tankarrange- ment was necessary for CG controland to limit the effect ofcrash condition pressureheads. The multibubbletanks ofConcept 3 retainedIsogrid stiffening of thetank walls. Thehoneycomb constructionactively cooled panels were retainedbut modified to be non-structural. A semi-structuralconfiguration similar to Concept 2 was consideredbut found to require complex support structure and resultin a rangedeficit. The majoradditional trade study conducted for Concept 3 involvedtank cross section optimization. Thistrade study involved tailoring the fuel tank within the elliptical fuselageto obtain maximum aircraftrange. Assessment of the range effect utilizedthe sensitivity curve of Figure 35 developedspecifically for Con- cept 3. This was accomplishedby comparing structural containment efficiency (weightof fuel/weight of containment structure) and volumetric efficiencies forseveral combinations of bubbletank configurations as shown inFigure 36. The results of thosecomparisons indicated that the five bubble cross section is bestin terms of structuralcontainment efficiency. Aircraft range levels

Legend: -Range Mrn (NM) - - TOGW Mg (lo3 Ibrn) I I 1 I I 175 180 185 190 195 200 O.W.E. - Mg I I I I I I 390 400 410 420 430 440 O.W.E. - lo3 Ibrn FIGURE 35 RANGE SENSITIVITY, CONCEPT 3

46 falloff for anybubble number above five, as indicatedin Figure 36. The fivebubble configuration was selected on thebasis of being a fighterand more easilyfabricated configuration, than the seven bubble configuration.

FIGUREOFMERIT THREE BUBBLE CROSS-SECTIONAL 73% AREA UTILIZATION WEIGHT EFFICIENCY W EIGH T FUELWEIGHT 17.6 WEIGHT STRUCTURE I FABRICATIONCOSTLO w HIGH MODERATE

RANGE RANGE PENALTY PENALTY

100 NM Mm

NUMBER OF BUBBLES FIG'URE 36 TANK CROSS SECTION SELECTION StructuralAnalysis The structuralarrangement for Concept 3 is shown inFigure 37. The processof structural weight refinement previously described was alsofollowed for Concept 3. Designloads, developed for the Concept 3 aircraft, are presentedin Figures 38 and 39. a. FiniteElement Computer Model - The Concept 3 structural model is illustratedin Figure 40. 1016joint degrees of freedom are usedwith this model.The resizingroutine described previously was also employed forthe Concept 3 analysis. b. Tank - Detailedanalysis of theConcept 3 tank was conductedin a manner similar tothat described for Concept 2. However in thisanalysis the fuselagecovering was considered to be completely non-structural. This analysisresulted in a tankweight of 14.54 Mg (32,047lbm) compared with the

47 Note: Active cooled panels cover all external surface except nacelles. r FuselageCover ActivelyCooled Panels // I

Speedbrake v Aft Transition Truss Links7 -

r Fwd Fuselage

A\ -Wing AttachLinks

Links

\ LPassengerCabin

Nacelle Module FIGURE 37 CONCEPT 3 STRUCTURAL ASSEMBLY BREAKDOWN 2.5 -

2.0 -

1.5 -

1.0 - LC 0.5 - Z In 2 0-L 7 m mL a 2 -0.5 - ~n v) -1.0 -

-1.5 - -2.0 - -2.5

Fuselage Station - meters L I I I I 1 I I I 0 40 80 120 160 200 240 280 320 Fuselage Station - ft FIGURE 38 CONCEPT 3 Net Aircraft Shear and Moment 2.59 Flight Condition

3.0 8 Ultimate Loads 2.0 - 6 Shear 4 1.0 - \ 2 LC \ 0 m -2

L -4 m -6 v) -6 -8 -3.0 - nz = 2.09 -1 0 -8 DW = 296.29 Mg -1 2 -4.0 - (653,204Ibm) '1 -10 -14 -5.0 -1 6 0 10 20 40 30 50 60 70 80 90 100 Fuselage Station - meters L I I I I I I I 1 0 40 80 120 160 200 240 280 320 Fuselage Station - ft FIGURE 39 CONCEPT 3 Net Aircraft Shear and Moment 29 Taxi Condition

I-

initial estimate of 10.57 Mg (23,292lbm). A large amount of the weight increasein this tank structure is attributableto the rings which act asthe wingcarrythrough-. A slight amountof weight, 213 kg (470 lbm) was. added to thetank for fatigue considerations in thetank rings and no weight addition was requiredfor fracture mechanics.

Cooling Sys tem Analysis The proceduresused to analyze the Concept 3 coolingsystem were similar tothose described for Concept 1. However,Concept 3 is a smaller aircraft and its moldlinecontours are different.These differences resulted in lower airframeheat loads and coolant flowrate requirements. Due tothe signifi- cantlydifferent fuselage shaping and volume utilization,the primary cooling systemcomponents were relocated to an areaforward of thetankage. This non- centralized location resulted in a weightpenalty due to larger distribution lines.These results are summarized in Figure 41. Similar tothe other air craftconcepts, the engine fuel flowrates during cruise are insufficientfor coolingthe entire airframe. It was estimatedthat the fuel heat sink caps- cityfor Concept 3 is alsoapproximately 50% ofthat required.

Hydrogen Fuel Tankage Insulation Weight = 3.1 1 Mg (6.855 Ibm) ThermalProtection System f Fuel Boiloff = 2.59 Mg (5,713 Ibm)

QVerficaltail = 7.6 MW (7.24 x lo3 Btu/sec) mvertical = 54 kg/s (1 19 Ibrn/sec)

Air Gap 39.7 MW (3.76 x 104 Btu/sec) 1.80cm (0.71 in.) Q~~~~~ mwings = 310 kg/s (681 Ibm/sec) Insulation '. 4.57 cm (1.80 in.)

Environmental Control System and Purge System Components Active Cooling System Heat Exchanger Active Cooling SystemWeight Component: Mg Qfuselage = 42.9 MW (4.07 x lo4 Btu/sec) Residual Coolant 8.92 mfuselage = 349 kg/s (769 Ibm/sec) Distribution Lines, etc. 1.35 Heat Exchanger 0.95 Pumps and Pump Fuel Req 2.51 Total 13.73

Note: Totals include subsystem requirements.

FIGURE 41 THERMODYNAMIC SUMMARY, CONCEPT 3

SECTION 8 FINALAIRCRAFT CHARACTERISTICS

The tradeoffstudies and design refinement process resulted in final designand performance characteris tics ofthe three concepts studied. These results are summarized in Figure 42. The lowestweight and best range is achievedwith the blended body integral tank aircraft, Concept 3. Group weight statementsfor each aircraft are presentedin Figure 43. Analyses were conductedto determine the relative producibility and ser- viceabilityof the three study aircraft. Although these analyses were greatly simplified,their relative values are believedto be accurate. Each analysis. concentrated on those areas wherestructural and arrangement differences were significant. Producibilityanalyses, summarized in Figure 44, includedboth material andlabor. Differing items such as thecenter fuselage cover, tanks and wing attachment were consideredin some depthwhile common items, such as thefor- wardfuselage, wing and tail, were includedin a productbilityfactor which CONCEPT I NON-INTEGRAL TANKS WING BODY O.W.E. = 190.2 Mg (419,234 LBM) RANGE = 8.69 Mm (4,690 NM) COMMON COST FACTORS: CHARACTERISTICS PRODUClBlLlTY1.0 109.9 rn SERVICEABILITY 1.0 t-" "-+ e WTFUEL = 108.9 Mg (240,000 LBM) (360.5 FT) e PAYLOAD = 21.8 Mg (48,000 LBM) CONCEPT 2 INTEGRAL TANKS (200PASSENGERS) WING BODY O.W.E. = 190.6 Mg (420,252 LBM) ENGINES (4) GE5/JZ6-C RANGE = 8.73 Mrn (4,715 NM) HOT NACE'LLE STRUCTURE COST FACTORS: , PRODUClBlLlTY 3.5 e CRUISE SPEED M = 6.0 I"-- 109.9m + SERVICEABILITY 1.2 e SURFACE ACTIVELY COOLED (360.5 FT) TO 394 K (25OOF) MAXIMUM CONCEPT 3 e T/W 0.55 (TAKE OFF) INTEGRAL TANKS = BLENDED BODY e W/S O.W.E. = 187.2 Mg (412,816 LBM) = 2.87 kPa (60 LBM/FT2) e fl RANGE = 9.20 Mrn (4,968 NM)

(328.5 FT) FIGURE 42 FINAL AIRCRAFT CHARACTERISTICS Concept 1 Concept 2 T Concept 3 Ms (Ibm) Ms (Ibm) Ms

I Structure A. Fuselage 1. Fwd 12.16 ( 26,800) 12.16 ( 26,800) 12.66 ( 27,900) 2. Center (Includes Fuel Tanks 29.39 ( 64,800) 29.98 ( 66,100) 32.25 ( 71,100) 3. Aft 1.72 ( 3,800) 1.72 ( 3.800) 1.91 ( 4,200) B. Remaining Structure 62.87 (138.600) 62.73 (138,300) 57.88 (127,600) II Propulsion Group 27.76 ( 61,200) 27.76 ( 61,200) 27.76 ( 61,200) III Systems A. Coolant Distribution System 15.15 ( 33,400) 15.20 ( 33,500) 13.74 ( 30,300) B. Remaining Systems 15.42 ( 34,000) 15.42 ( 34,000) 15.37 ( 33,900) IZL UsefulLoad 25.67 ( 56,600) 25.67 ( 56,600) 25.67 ( 56,600) ll O.W.E. 190.14 (41 9,200) 190.64 (420,300) 187.24 (41 2,800) 3ZI Fuel 108.86 (240,000) 108.86 (240,000) 108.86 (240,000) Usable ,106.27 (234,300) 106.30 i234.400) 106.27 (234.300) Boil-off 2.59 ( 5,700) 2.56 ( 5,600) 2.59 ( 5,700) Yll TOGW 299.0 (659,200) 299.5 (660,300) 296.1 (652,800)

FIGURE 43 FUSELAGE/TANK-GROUP WEIGHT STATEMENTS

Item Concept 1 Concept 2 1 Concept 3

Welding 1 7 Forming 1 2.5 I 2.5 Material 1 5 3 Machining FuselageFrames and Bulkheads 1 2 0.4 0 Tank to Fuselage Ties 1 15 32 TankFrames 1 24 15 0 TankWall 1 31 35 Tank Ends 1 9 9 Wing Attachment 1 1 0.1 Overall Machining 1 20 15 Assembly 1 3 5 Center Fuselage" 1 10 8

Total Vehicle Cost 3 1 3.5

'Includes tank, wing supports, and fore and aft stress links FIGURE 44 RELATIVE COST RATIOS

54 was common to all threeaircraft. Relative cost ratios are presentedfor the fuselage/tank area and the total aircraft andprovide a measure of the relative fly-awaycost for each concept. Since all threeaircraft concepts carried the same fuelload andnumber of passengers, an indication of therelative opera- tional costs may beattained bycomparing aircraft range as well as theservice- abilityfactors discussed below. These ratios show Concept 1 tohave the lowest productioncost. A similar study for Concepts 2 and 3, replacingintegrally stiffened tank walls with plain skin monocoque walls, resulted in reduction of totalvehicle cost factors from 3.5 to 1.6 forConcept 2 andfrom 3 to 1.8 for Concept 3. Serviceabilityanalysis also concentrated on thosedifferences in aircraft arrangementwhich significantly affected maintenance actions. The netresult of theseanalyses given in Figure 45 indicatedConcept 1 to bethe most easily maintainedwith Concepts 2 and 3 being more difficult by factors of 1.2 and 1.3 respectively.

~ ~ ~_ ~ T Concept Service or General Maintenance Action ~~ - 1 2 3 .. . . ~- ~~~~ Structural Tank Repairs 1.oo 1.30 1.40 Actively Cooled Panel Leak Inspection 1.oo 1.20 1.40 Actively Cooled Panel Removal 1.30 1.20 1.oo Actively Cooled Panel Manifolds and Controls 1.oo 1.30 1.40 Link and Lug AdjudRepair 1.oo 1.60 1.80 Coolant Supply Lines 1.oo 1.40 1.50 Coolant Return Lines 1.oo 1.40 1.50 Heat Exchanger Unit 1.oo 1.30 1.40 Nitrogen Purge System 1.40 1.30 1.oo FuelFeed Lines 1.oo 1.30 1.50 Fuel Boost Pumps 1.oo 1.20 1.40 Fuel Transfer Controls 1.oo 1.30 1.30 Plumbing Repairs 1.oo 1.20 1.30 ElectricalRepairs 1.oo 1.20 1.30 Flight Control Cables 1.oo 1.20 1.40 Average Level of Difficulty 1.04 1.28 1.37 Normalized Level of Difficdty 1 1.2 1.3 ~ ~-~~~ ~ ~~ ... Comparative Ratings (Degree of Difficulty) 1.0 = Concept with Lowest Mean Time to Complete Maintenance Action iUsed as Baseline) 1.5 = 50% Greater Time to Complete Action Compared to Baseline 1.8 = 80% Greater Time to Complete Action Compared to Baseline

FIGURE 45 RELATIVE SERVICEABILITY FACTORS

SECTION 9 OBSERVATIONS AND CONCLUSIONS Thisreport has presented the results from a structuraldesign study of activelycooled, hydrogen fueled, hypersonic transport aircraft. In addition to the final design and performance description of the three aircraft con- ceptsstudied, detailed design trade studies and sensitivity studies were performedin the early phases. Simplified maintainability and producibility studies were alsoconducted. Thus, a broadbasis is availablefor assessing theimportant design factors which must be considered for this class of aircraft . The finalaircraft characteristics, weights, and relative costand ser- viceabilityfactors were presentedin Section 8. From Figures 42 and 43, it is seenthat Concept 3, theblended body, integral tank aircraft has the lightestweight and the greatest range capability, (over 0.47 Mm (250 NM) more thanthe others). This superior performance is a result of thebetter aerodynamiccharacteristics and higher volumetric efficiency. The structuralweight differences between integral and non-integral tankage are small. This was examined in some detailin order to understand thefactors which influence this finding. Figure 46 shows a typicalcross section through the fuselageltank structurefor each aircraft. In all cases thecooled wall is made ofaluminum honeycomb panelswith "dee" shaped cooling passages bonded to the outer skin. Thesepanels absorb the aerodynamic heating. The tank wall, whichserves the function of fuelcontainment and pressure vessel, is monocoque structure for thenon-integral concept and isogrid stiffened structure for the integral Concepts. The outer wall coolingfunction and tank wall pressurefunction are thus common to all concepts.However, for each concept, the fuselage bendingfunction is performeddifferently and wing loads are carried differently. Inthe non-integral Concept 1, thecooled outer wall servesthe addi- tionalfunction of providing strength for the fuselage primary structure. This is accomplishedby increasing the depth of the honeycomb paneland the thicknessof the outer skin. However, theserequired increases are small, so thecooled outer wall panels are onlyslightly heavier than those of the integral concepts.

57 CONCEPT 1 CONCEPT 2 0.6 mm - 1.1 mrn- 1 0.6 mrn 7 f =I 0.4 rnm f

15.2 crn FRAMES AT 0.91 m ISOGRID TANK WALL AND STRUCTURE t = 1.5mrn. 1.8mm (i = 2mm - 2.8rnrnl I 2.5 cm t MONOCOQUEWALLTANK 30.5 cm FRAMES AT 0.91 m io 2.74 m 2 mrn- 1 CONCEPT 3 0,6m-

TANK PRESSURE = 138kPa

Note: Not to Scale ISOGRID TANK WALL AND STRUCTURE A i = 1.5rnm - 1.8rnrn li = 1.8rnm - 2.8mm) 2.5 crn 223 cm FRAMES AT 0.91 m to 2.74 m

FIGURE 46(a) STRUCTURAL ARRANGEMENTS S.I. Units CONCEPT 1 0.025.0.045 in.-

8.13 in. 6 IN. FRAMES AT 3 Fl ISOGRID TANK WALL AND STRUCTURE t = 0.060-0.070In. li = 0.081-0.1 10in.) 0.98 in. 12 IN. FRAMESAT 3 TO 9 FT

0 TANK PRESSURE = 20 PSlG

Note: Not to Scale ISOGRID TANK WALL AND STRUCTURE t = 0.058-0.070in. ii=0.070-0.llOin.~ 0.98 in. U 9 IN. FRAMES AT 3 TO 9 Fl

FIGURE 46(b) STRUCTURAL ARRANGEMENTS Customary Units

58 Forthe integral tank concepts the tank wall providesfuselage primary structurestrength using a stiffenedskin approach. Of thestiffened design candidatesanalyzed, an is,ogrid stiffening pattern produced the lightest weight.If the volume of the stiffening members is averagedover the surface, theequivalent weight thickness (t ofFigure 46) is onlyslightly greater thanthat for the non-integral monocoque tank wall. Thus,the stiffened tank wall is only slightly heavier than the unstiffened non-integral tank wall. The combination of outer panel weight plus tank weight results in approx- imatelythe same structuralweight for each of the three aircraft concepts. Thisresult appears to be strongly influenced by the fact that actively cooledstructure is used. While notproven by this study it appearsthat use of activelycooled st.ructurefavors the non-integral structural concepts. Consider the complete structuralsystem of the non-integral Concept 1. It essentiallyhas two structuralelements. The outerelement serves as a heatshield as well as primarystructure. The innerelement contains the fuel. The integral Con- cepts 2 and 3 alsohave two structuralelements. The outerelement serves as a heatshield and the inner element serves the dual functions of primary structureand fuel containment. Consider the situation if a non-actively coQledstructure were employed,say a radiation heat shield/insulation system. Then thenon-integral Concept 1 would requirethree structural elements:an outer heat shield/insulation, a primary structureand an inner fuelcontainer. However, theintegral Concepts 2 and 3 would still only require two structuralelements. Thus it appears logicalthat the addition of a thirdstructural element to the non-integral concepts would result in anincremental weight penalty which is notpresent for the actively cooled structure. The small differencesin structural weight are alsoinfluenced by the typeof structure selected, the tank design pressures, the material properties ofthe tanks at cryogenictemperature and the very large size of the circularfuselage. Change to anyor all ofthese elements could change the results.For example, the large size of the vehicle is a dominantfactor in allowingthe usage of monocoque structurefor the non-integral Concept 1. If thetank diameter were cut in half, the wall thicknessneeded to sustain pressurewould also be halved. The equivalentbending strength would, however, bereduced by a factorof eight. To regainthe necessary bending strength,substantial stiffening would be required. Thus, tank pressure and size are extremelysignificant elements in driving the comparative results. The principal aircraft performance results are plotted on Figure 47, to provide a quick visual comparison of the three aircraft concepts studied. Consideringthe effect of tankconstruction, it is seenthat there is very little differencebetween the circular body non-integral(Concept 1) and circular body integral (Concept 2) aircraft. Each resultsin comparable rangeand TOGW. However, when consideringthe effect ofbody shapethere is a clear advantagefor the integral tank blended body (Concept 3) overthe integraltank circular body(Concept 2). As statedpreviously, this is due to the improvedaerodynamic characteristics and better volumetric utilization. Tc provide a basis ofcomparison between Concepts 2 and 3, it is seen that by offloading 4.54 Mg (10,000 lbm)of fuel,the Concept 3 rangebecomes the same as Concept 2, but TOGW is significantlyreduced. In summary, theresults of this study are judgedto have identified significanttechnical factors relating to integral and non-integral tank designs, in addition to providing performance effects of body shape.

60 5000 - BLENDED BODY - INTEGRAL 9.2 WF = 108.9 Mg (240,000 Ibm) O.W.E. = 187:2 Mg (412.800

4900 -

RANGE RANGE - 4800 Mm NM

4700 -

CIRCULAR'. NON4NTEGRAL- WF = 108.9 Mg (240,000 Ibm) O.W.E. -190.2 Mg (419.2001bm) 4600 - 8.5 I I 29 0 295 300 295 8.61"v290 305 TOGW - Mg

640 650 660 6.10 TAKEOFF GROSS WEIGHT - 1000 LBM

FIGURE 47 BEST PERFORMANCE WITH BLENDED BODY

61 The principal overall study conclusions are: a. Integralvs Non-Integral Tanks 1. There is very little weightdifference for circular body concepts. 2. Producibilityand maintainability factors favor the non- integral tank concepts. 3. Multi-bubbleconfigurations require more welding and assembly time andthus are more costlyto produce than circular tank configurations. b.Circular vs Blended Body Shape 1. Higher LID andgreater volumetric efficiency are achievable with the blended body. 2. Better performance(i.e., greater range for the same weight) can be achieved with the blended body shape. c. VolumetricEfficiency 1. The efficientutilization of all availablevolume is a dominant factorin maximizing performance. This was shown in a number oftrade studies as well as inthe comparison of the circular andblended body shape. The studies of passengercompartment arrangementand location, along with studies of dome shapeand number oftanks, all showed significantperformance improve- ments when volumetric utilization was increased. 2. Designcompromises which increase fuel volume can improve aircraftrange even though the compromise results in increased structuralweight. Therange sensitivity curves for Concept 1 and 2 show thatrange will be increased if for every additional unit of fuelweight less than 1.6 units of structuralweight areadded. Similar valuesfor Concept 3 are 1.9units of structuralweight per unit of fuelweight. Thus the importance of achievinghigh fuel volumetric efficiency is againevident. d. StructuralDesign 1. Monocoque tank walls providean attractive fabricationapproach from a producibilitystandpoint. However, they may resultin increasedweight and decreased range, depending upon mission requirements.For Concept 2, useof a monocoque ratherthan

62 stiffened tank wall results in a 71% increase in tank weight and a rangeloss of .about 463 km (250 NM). ForConcept 3, the figures are 16% and 185 km (100 NM) . 2. The monolithicnon-integral tanks were designedto burst pressurerequirements and not limited by fatigue or fracture mechanicsconsiderations. However, the stiffened integral tanks were critical at the tank frames for fatigue design, althoughno weight was addedfor fracture mechanics. Thus forthe design service life (10,000 hr)used in this study fatigueand fracture mechanics requirements generally had negligibleperformance effects. However, for a higherservice liferequirement these considerations would be more significant. 3. Vehiclesize and tank design pressures had a stronginfluence on theresults. Thecombination of these two factorsallowed the non-integral tank to be of monocoque designand still be at minimum weight.If tank pressures and/or the vehicle size were reduced, the minimum weightdesign would probably have been a stiffened structure design. 4. Provisionsfor gaseous N2 purgingbetween the panels and the tankagecan be met with minimal structural weight addition. e. ThermalDesign 1. The enginefuel flow rate demands duringcruise are 52% to 62% ofthose that would be required if the entire airplane surface exclusiveof the nacelle were cooled. 2. The useof internal insulation in the tanks was uncompetitive, becauseof GH2 permeationof the insulation. Covering the insulation with a gaseoushydrogen vapor barrier would have distinctadvantages. Thermal protection weight and volume requirementswould be reduced and structural design problems wouldbe minimized without the need for thermal expansion allowances. 3. Trajectorytailoring to minimize peak transient heating loads is importantin minimizing the size and weight of the cooling systemcomponents and the coolant distribution lines.

63 4. As was foundin the nacelle cooling study, designing the regions subjected to high heating rates as hot structure may reduce air- craftweight. It appearsthat this observation would alsoapply to remote areas, whereactive cooling would drive up the size andweight of thecoolant distribution system. f. Fly-awayCost 1. Producibilityfactors favor the simple monocoque non-integral tankapproach. g. OperationalCost 1. Maintainabilityfactors favor the non-integral concept. 2. Comparison of thethree concepts at equalrange gives an indication of comparativefuel costs, since the payload was the same forall three concepts. This comparison favors the blended body shapeover the circular body shape,since slightly less fuel is used for the sane mission.

64 SECTION 10 CRITICAL AREAS REQUIRING ADDITIONAL RESEARCH AND TECHNOLOGY

Vapor Barriers Theaccommodation of thermal strains and the requirement for integrally machined stiffening concepts for the tank walls are themajor contributors tothe high relative costsof integral tankage. The development of a viable vapor barrier against gaseous hydrogen leakage would allow the practical use ofinsulation on theinside of the tank wall. Thisin turn wouldreduce or eliminate the temperature differential between the tank wall and actively cooledstructure, permitting use of lighter and simpler structural concepts. Largestructural thermal deflections would not have to be accommodatedand volumetricefficiency would be enhanced. It is possiblethat the honeycomb constructionactive cooling concept and load carrying tank wall couldbe combined in a manner toaccomplish the load carrying function as well as fuelcontainment function. Further,for the structural approaches presented in this report, vapor barriers would resultin undegraded insulation properties and hence volumetric efficiencywould be enhanced and weight lowered due to reduced insula- tionrequirements. Vapor barrierresearch and development therefore offers interesting designoptions.

-______Materials There is currentlyconsiderable interest in the use of cryogenic hydrogen foraircraft systems, particularly as anenergy conservation measure. Practical construction of tankage for this fuel will dependon extensive new knowledge of material propertiesspecifically suited for this purpose. The structuraldesigner needs material propertiesand design allowables for cryo- genictankage construction materials. Such tankstructures probably will be weldments.Thus welding data, along with fatigue and fracture mechanics data, are needed.Development of light alloy weldable materials would enhancethe ability to design reliable lightweight tankage for future hydrogen fueled aircraft.

65 RadiativeThermal Protection Systems Studieswhich match airframe cooling requirements to available fuel heat sinkcapacity are needed toestablish realistic systemweights. The potential ofcombining insulation and active cooling in panel designs should be evaluated.Radiative systems will probablybe required over part or all of the surfaces to reduce the total heat load absorbed by thecooling system and match theavailable heat sink capacity. Suchsystems would offer the potential of failsafe capability and also could even reduce the total system weight. It appearsthat the design and development of radiativeprotection systems will beneeded for actively ccoled aircraft.

Subcooled LH Fuel "- "- 2 The effect of thetank size and design pressures on establishing the tank wall thicknessfor Concept 1 was dominant inthe study. Use ofsub- cooled(or slush) hydrogen would allow a lowerfuel system tank pressurization and also reduce the amountof boil-off fuel lost during the mission (bothground operations and flight). Fuel system components for handling sub-cooledhydrogen do not represent major technological advancements. However, thepotential benefits should be further investigated to determine the optimum degreeof sub-cooling in order that design requirements can be established.Based on these requirements, system design and development investigations wouldbe of value.

CoolingSystem Optimization Investigationsof several methods that could result in active cooling systemweight reductions are warranted.Trade-off studies involving the followingconsiderations are suggested: a. Reducecoolant flowrate requirements by maximizing allowable surface temperaturesand outer skin thermal gradients. - Whilestructural weight penalties may result,the tradeoff with cooling system weight should be clearlyestablished. For example, designing for a 422 K (300°F) maximum temperaturerather than 394 K (250°F)would reduce flowrate requirements by nearly 40%. Increasingthe allowable skin thermal gradient from 56 K (100°F) to 72 K (.130"F)would reduce the number of coolant tubes required by about 10%.

66 b.Optimize coolant system design pressure. - A systemdesign pressure of 1.03 MPa (150 psi)absolute was chosenfor this study. A higherdesign pressurewould permit larger pressure drops in the distribution lines, hence smaller lines and less residualcoolant. Obviously, higher design pressures necessitatestructural weight increases and increased pumpingpower. It is estinated that a designpressure between 1.38 MPa (200psi) absolute and 1.72 MPa (250psi) absolute, would reduce system weight by about 907 kg (2,000 lbm) . c. Establishthe significance of centralizing the location of major coolingsystem components. As discussedin Section 7, theheat exchanger locationfor Concept 3 resultedin a significantweight penalty. The penalty involvedin providing volume for components at a morefavorable location shouldbe assessed. d.Refine feeder line sizing technique. - Eachfeeder line could be sizedto match the local available pressure drop between the main supply and returnlines. This study considered only a constantpressure drop per unit linelength in establishing feeder line sizes. It is estimatedthat, by consideringlocally higher line pressure drops, a weightsavings in the order of 454kg (1000 lbm) is possible. e. Consideralternate line routing schemes. - There are aninfinite number of possibleline routings. While no attempt was made duringthis studyto find an optimum configuration, it seems logicalthat benefits could bederived.

67 I

11. REFERENCES

1. James C. Ellison,"Investigation of the Aerodynamic Characteristicsof a HypersonicTransport Model at Mach Numbers to 6," NASA TN D-6191, 1971, (Unclassified). 2. T. Nobe, "A Fuselage/TankStructure Study for Actively Cooled Hypersonic CruiseVehicles - AircraftDesign Evaluation," NASA CR-132668, June1975. 3. James E. Stone, "A Fuselage/TankStructure Study for Actively Cooled HypersonicCruise Vehicles - Active CoolingSystem Analysis," NASA CR-132669, June1975. 4. Allen H. Baker, "A Fuselage/TankStructure Study for Actively Cooled HypersonicCruise Vehicles - Structural Analysis," NASA CR-132670, June1975.

NASA-Langley, 1976 CR-2651 69