Hypersonic cruise aircraft propulsion integration study, volume 1

NASA Contractor

CR-158926-1

Report

Hypersonic Cruise Aircraft

Propulsion Integration Study

Volume I

R. E. Morris

G. D. Brewer

LOCKHEED-CALIFORNIA COMPANY BURBANK, CALIFORNIA CONTRACT NASI-15057 September 1979

[ rL A

National Aeronau[:cs and Space AdmJnistrat;on Langley Research Center HampTon, Virginia 23665 FOREWOKD This is the final report of a study made under Contract NASI-15057 for NASA-Langley Research Center, Hampton, Virginia, Volume I includes the study guldellnes, the candidate configuration analysis and selection, propulsion concepts, final propulsion evaluation and comparison, and the study conclusions and recommendations.

Volume II presents supporting aerodynamic, propulsion, and weight technology data as well as the selected candidates conflgura_ion analysis and refinement of the final baseline vehlcle used for evaluation of the two propulsion concepts described in Volume I.

The Lockheed-Callfornla Company was the principal contractor to NASA and the work was perfomed in the Commercial Advanced Design Division at Burbank, California. The following individuals were the main contributors: Daniel Brewer, Study Manager Rober Morris, Project Engineer Jerry Rising, Aerodynamics Marvin Baxendale, Aerodynamics • Roger Jensen, Weights Chris Monoleos, Aircraft Synthesis Mr. Joe Watts of the Hypersonic Aerodynamics Branch of NASA - Langley served as Technical Monitor.

.6 TABLE OF CONTENTS Page FOREWORD ............................ i LIST OF FIGURES ......................... iv LIST OF TABLES ......................... v SUMMARY ............................ i SYMBOLS 4 . • , • • • • , • • , , • . , • , , , • , • , • $ • • • 1.

INTRODUCTION ................. • • • • • . • . . 7 2 • • , , • , . , • , 3.

3.1 3.1.1 Data acquisition and review 9 • , • • , , • . , , • , , • , 3.1.2 3.1.3 Configuration screening ................. ii 3.1.4 Vehicle synthesis .................... ii 3.1.5 3.1.6 3.2 Configuration Refinement ................ • . 20 3.3 3.3.1 3.3.2 4.

4.1 4.2 4.3 4.4 5.

5.1 5.1.1 5.1.2 5.1.3 5.1.4 Subsonic cruise range ................ -. . 47 5.2 5.3 il

TABLE OFCONTENTS (Continued)

Page

5.4

6.

STUDY CONCLUSIONS• • • " • • * • • • e • • • e • • • • • • • o 65

6.1

6.2

6.3

7.

REFERENCES iii LIST OF FIGURES

Figure

Page Work Plan, hypersonic cruise aircraft propulsion integration study .............................

i0 ASSET vehicle synthesis program schematic ...........

HYCAT-I general arrangement ..................

HYCAT-2 general arrangement ..................

HYCAT-2 cabin cross section ..................

HYCAT-3 general arrangement ..................

HYCAT-4 general arrangement ................ . .

HYCAT-4 propulsion installation ................

i0 II General arrangement, baseline version of HYCAT-IA .......

Common variable geometry inlet, turboJet-ramjetsystem schematic ". .........................

HYCAT-IA, final general arrangement .............

Mission climb history, turbojet-ramjet system ........

Installed thrust coefficient, ramjet with fixed diverter . . .

Installed specific impulse, ramjet with fixed diverter ....

Cruise part power performance, ramjet with fixed diverter . .

2O Installation drag comparison, TJ-RJ and TJ-SJ systems ....

Inlet mass flow comparison, TJ-RJ and TJ-SJ systems .....

Performance comparison, Math 3.5 to 5; TJ-RJ and TH-SJ Systems ...........................

iv LIST OF TABLES Page Table 6a Point Design ChJracteristics of HYCAT-IA at Ranges of 7408, 6b Point Design Characteristics of HYCAT-IA at Ranges of 4000, 7a Weight Build-Up Comparison of Turbojet-Ramjet System Installed 7b Weight Build-Up Comparison of Turbojet-Ramjet System Installed in Turbojet-Scramjet Point Design Aircraft (Customary Units). . 51 8a Comparison of Mission Fuel Consumption of Turbojet-Ramjet System Installed in Point Design Turbojet-Scramjet Aircraft 8b Comparison of Mission Fuel Consumption of Turbojet-Ramjet System Installed in Point Design Turbojet-Scramjet Aircraft Baseline Aircraft Cruise Comparison of the Turbojet-Ramjet 9a 9b Baseline Aircraft Cruise Comparison of the Turbojet-Ramjet ii Weight Fraction Comparison - Point Design, Scramjet and 12a Aircraft Weight Comparison of Turbojet - Ramjet System with 12b Aircraft Weight Comparison of Turbojet - Ramjet System with V HYPERSONIC CRUISE AfRCKAFT PROPULSION INTEGRATION STUDY R. E. Morris, and G. D. Brewer Lockheed-California Company Burbank, California SUMMARY This report, consisting of Volumes I and 11, describes the work done by the Lockheed-California Company on the NASA Hypersonic Cruise Aircraft Pro- pulsion Integration Study, Contract NASI-15057. The primary emphasis was to evolve the most promising conceptual vehicle and propulsion integration ap- proach for a liquid hydrogen fueled, Mach 6 transport capable of carrying 200 passengers 9 260 km (5 000 nm).

The work was conducted in two phases with the initial phase being a generation and screening of candidate vehicle configurations, comparative analysis of the two most promising concepts, selection and design refinement of the surviving candidate. The final phase used this selected configuration as the baseline aircraft in the comparative evaluation of two propulsion integration concepts: A turbojet engine with a retractable fnletused for takeoff, accelera- tion and landing, together with separate flxed-geometry dual-mode combustion scramjet engines for cruise (Turbojet-Scremjet System).

A turbojet engine with a separate variable-throat subsonic combustion ramjet engine with both engines obtaining air from a common variable- geometry inlet (Turbojet-Ramjet System).

Other trade studies included the effect on aircraft gross weight of such variables as wing geometry, field length, approach speed, range, propulsion installation drag, gross thrust vector angle, range capability during all subsonic cruise and growth sensitivity.

The major conclusions drawn from the initial or vehicle configuration selection and refinement phase are: • The gross weight of aircraft to perform the design mission are in the 272 160 to 362 880 kg (600 000 to 800 000 Ib) class.

• The lift provided by a flattened fuselage forebody is important in improving hypersonic L/D and in providing the flow field and geometric J width necessary for the propulsion installation. This is of particular importance in hydrogen-fueled aircraft with a large potential fuselage to wing planform a_ea ratio.

The use of a horizontal tail in the selected configuration was required for trim purposes and provided a favorable tradeoff by allowing the use of drooped ailerons to obtain more low speed lift with the final payoff being the reduction of wing size and weight.

A further benefit is the reduction of the neutral point variation with Math number.

• The most critical design criterion is to meet the landing field length constraint without increasing the wing aspect ratio or reducing the wing loading, both of which options result in increased gross weights.

The propulsion system should be integrated with the fuselage to avoid • excessive wave and friction drag. It should also be located far enough forward for balance purposes and to allow for takeoff rotation without requiring a long main gear for clearance. Further benefits are the reduction of propulsion moments when the system is located near the center of gravity, and a reduction in the boundary layer displacement thickness. Adverse effects of the fuselage boundary layer could dictate the use of wing-mounted propulsion nacelles.

• The location and optimum inclination of the gross thrust vector can make a significant reduction in cruise fuel flow by reducing the aerodynamic lift required and consequently the drag.

• Based on supersonic transport design experience and the high growth sensitivity of the hypersonic transport, the imposition of airport noise constraints would have a very adverse impact on vehicle size although it is possible that this could be mitigated to some extent by a variable cycle accelerator engine in which, as a secondary benefit, the subsonic SFC could be improved thereby reducing the reserve fuel consumption.

The results of the final propulsion integration study phase indicate that to perform the design mission, the vehicle using the turboJet-scramjet system would require a gross weight of approximately 351 000 kg (774 006 Ib) compared to 278 000 kg (613 000 Ib) for the turbojet-ramjet propulsion system.

In each case the aircraft was optimized with respect to wing loading, thrust to weight and capture area or cowl size while meeting the critical perfor- mance constraints. Both aircraft flew the same mission and had the same reserve fuel requirement in subsonic flight. The major conclusion from this phase is that the difference in gross weights are due, not to the engine com- bustion mode (subsonic vs supersonic), but to the following: • The reduction in both mission fuel consumption and installed propul- sion weight mode possible By the use of a common variable geometry

inlet for both the turbojet and ramjet engines. The reduction in

spillage drag of the common inlet in the critical transonic region allows a smaller cowl size and reduced fuel consumption both in acceleration and subsonic cruise.

The use of this variable geometry inlet increased the inlet air flow (and thrust) in the critical Mach 3.5 to 5 region after turbojet shutdown.

The net result is that the turbojet-scramjet system is penalized in both fuel consumption and installed weight caused by high subsonic/transonic spillage drag and by low thrust in the Mach 3.5 to 5 region due to a lower mass flow resulting from the fixed geometry scramjet engine.

The primary recommendation, considering the propulsion application to a transport mission, is to pursue the use of a common inlet for the acceleration and cruise engines and to provide a higher thrust level in the Mach 3 to 5 region by variable goemetry or other means.

The majority of the remaining recommendations were the result of uncer- tainties in the prediction methods used in the study. Testing and analytical correlation is required in the following areas: Demonstrate that either the variable or fixed geometry engines (inlet + combustor + nozzle) could operate efficiently while ingesting the boundary layer from the long fuselage _orebody.

If a diverter is required for either system what is the low speed drag and what lift contribution is caused by the shock field impingment on the fuselage or wing underside?

• Determine by test the spillage lift and drag forces in the transonic r_gion.

Simulate propulsion flows to determine base drags and moments.

• Further work is required to define the comparative weights and cooling requirements of both propulsion systems.

SYMBOLS SI Customary Units Units

D A

area ft 2 m flow field streamtube areas A_o, A o , ft 2 m A I Z A._ minimum inlet area ft 2 m #.

ramjet inlet area m ft 2 A 3 geometric exit area ft 2 A 6 m A inlet geometric capture area m ft 2 C A exhaust flow area ft 2 m ex APU auxiliary power unit Aft aspect ratio ASSET Advanced Systems Synthesis and Evaluation Technique - Lockheed computer program C chord m ft mean aerodynamic chord m ft drag coefficient C D D drag lh kg FAR Federal Aviation Regulation net installed thrust F N N ib net uninstalled thrust FNj N Ib net sea level static thrust N FNsl s ib HYCAT Hypersonic Cruise Aircraft Technology IOC initial operational capability ISP specific impulse Ns/kg $ec Keas knots equivalent airspeed m/s kts SI Customary Units Units

L length, aerodynamic lift

LE leading edge

n/D lift to drag ratio

liquid hydrogen LH 2 math number M m ft }b_C mean aerodynamic chord free stream math number M_o flow field local m,:ch numbers M 0, M 1 , M 2 kg Ib OEW operating empty weight Pa lb/in 2 P _tatic pressure Pa ib/in 2 total pressure PT Pa lb/ft 2 dynamic pressure q RJ ramjet m ft2 wing reference area S, SRE F kg/hr/daN ibm/hr/Ib SFC specific fuel consumption SJ scramjet (supersonic combustion scramjet) sea level static SI.S, sls wing thickness ratio

T/C, tic

TJ turbojet sea level static thrust to aircraft gross T/W, FSLs/W 4N/kg weight kg Ib W, Wg gross weight kg/m 2 ib/ft 2

w/s wing loading

rad deg angle of attack SI Customary Units l;nics angle of gross thrust rad deg _0", 61 boundary layer displacement thickness m ft A sweep angle rad deg fuel-air equivalence rstio D !

.° 1. INTRODUCTION Thin is the Volume I final report of a study performed by Lockheed- California Company for the Hypersonics Branch of NASA-Langley Research Center.

The primary purpose of the work was to evolve the most satisfactory conceptual vehicle configuration and propulsion integration approach for a Mach 6 trans- port aircraft capable of carrying 200 passengers 9260 km (5000 n.mi.).

Hypersonic aircraft of the future will require propulsion systems which operate in two modes; one mode for takeoff, landing, and acceleration through the subsonic/supersonic speed regime and another mode for acceleration and cruise at Mach numbers above about 3.5. Many of the characteristics and requirements of the hypersonic cruise mode are not compatible with subsonic operation and many of the characteristics of the subsonic mode are not com- patible with the hypersonic speed regime. Considerable ingenuity and effort will be required to achieve a total system which circumvents the potentially high off-design performance penalties of either system.

Past studies of hypersonic cruise aircraft have not dealt in depth with the subsonic and transonic performance problems of hypersonic configurations; consequently the study effort was directed at the integration of the subsonic/ supersonlc/hypersonic propulsion systems with the aerodynamic design of the airframe.

In the first part of the study numerous configuration design approaches were considered. Some were rejected almost immediately for obvious reasons in _pite of their offering some unique advantage which led to their being st_gested in the first place.

Those aircraft and propulsion configurations which ap_=ared to be gener- ally promising were sized and design layout drawings were made. These concepts were screened qualitatively, then selected designs were evaluated quantita- tively using the Lockheed proprietary vehicle synthesis computer program, ASSET.

The results of the vehicle screening evaluation were used to select a preferred aircraft design concept for a more detailed propulsion integration concept analysis in the final effort reported in this volumn.

Vol II contains supporting data including an 6xplanation of technical methods which were used and configuration details which'were significant in the evaluation of the final vehicle concept.

2. STUDY CUIDELINES

The choice of a commercial transport to represent the mission to serve

as a basis for a design study of hypersonic aircraft was an arbitrary one, but to ensure consistent criteria for comparison purposes the following guidelines similar to current practice were used: i. Design mission: 200 passengers - 9260 km (5000 n.mi.) range -Mach 6 cruise.

Accommodations comparable with current supersonic transport concepts.

// 2. IOC date: 2000. Consistent advanced aircraft technologies were used.

3. Performance and environmental constraints congistent with practices at current large international airports. The performance at low speeds must be compatible with the airport aids and other aircraft in the airport environment. For example: ¢ Speed in controlled airspace 128 m/s (250 keas) maximum • Minimum engine-out climb gradient z 0.030 • Maximum FAR field length = 3200m (i0 500 ft) 4.

LH 2 assumed available at all airports.

5.

Requirements of FAR 25 (airworthiness standards) to be met where applicable.

6.

As a design goal, the aircraft llfe to be commensurate with current aircraft.

7.

The primary evaluation criterion used in selecting preferred designs was minimum takeoff gross weight.

Design allowances and requirements for the mission included the following: • An allowance of i0 minutes at ground idle power provided for taxi out and taxi in.

• One mhlute at maximum power provided for takeoff.

• Maximum speed below 3048m (I0 000 ft) to be 128 m/s (250 kias).

• Six minutes air maneuver time for landing.

• Fuel reserves: 5% of block fuel plus subsonic flight at optimum altitude and speed to a 482 km (260 n.mi.) alternate airport, plus 30 minutes loiter at 4572m (15 000 ft).

• Descent to be at equilibrium glide (L/D maximum). Turbojets to be turned on at Mach .8 at flight idle power to provide hydraulic and electric power. This power is supplied by an M'U when the turbojets are not running.

3. 'rECtlNICAI+ APPROACII in accordance with tile objective of developing a preferred configuration for a hypersonic transport aircraft, the initial phase of the study was aimed .it explorh_g all feasible concepts. The final phase involved a more detailed dt+sign study of propulsion concepts in a defined configuration selected as a resuit ,+',t: the screening analysts.

3.1 Candidate Configuration Analysis and Selection The study plan is graphically illustrated in figure 1.

3.!.1 Data acquisition and review. - in view of the basic requirement for a morphological a0proach to consider all feasible aircraft configurations, the first step in the process was to obtain information about previous design studies and to review the conclusions which had been reached concerning each.

In addition, the latest information which could be obtained about turbojet and turbofan engines that might be used for takeoff and acceleration to .Math 3.5, and on dual-mode convertible scramjet engines that were suitable for operation from Math 1.0 to Math 6.0 was explored.

A study by Lockheed (reference 1) was useful in providing realistic size, weight, and design requirement information about tile aircraft Ltl 2 fuel system .lilt] its major components.

Th is rev iew of pertinent data on hypersonic vehicles propulsion and hydrogen technology was used in the generation of candidate aircraft conf igurat ions.

3.1.2 Aircraft configuration conceptuali::ation. - As many aircraft de- sign cot:eepts as possible were postttlated during tile study. Any configuration which appeared to offer merit was considered. Innovative ideas were onCOLI raged.

Q CONFIGURATION SCREENING DATA ACQUISITION AND AIRCRAFT CONFIGURATION EVALUATE ALL SUGGESTED REVIEW CONCEPTUALIZATION CANDIDATE AIRCRAFT CONFIGURATIONS: SOURCES: CONSIDER INNOVATIVE • INITIAL SCREENING: ARRANGEMENTS AND • INDUSTRY QUALITATIVE MODIFICATIONS OF • NASA ASSESSMENT OF • AIR FORCE • LIFTING BODY ALL CANDIDATES • WINGED BODY SUBJECTS: DESIGN EVALUATION • BLENDED BODY OF SELECTED • VEHICLE CONFIGURATION STUDIES ESTABLISH CANDIDATE CANDIDATES • TURBOJET/TURBOFAN DESIGNS AIRCRAFT • SCRAM JET PROPULSION • FINAL SCREENING CONFIGUP:4TIONS • PROPULSION INTEGRATION QUANTITATIVE EVALUATION OF TWO PREFERRED CANDIDATES n PROPULSION CONCEPTS MID-TERM ORAL EVALUATION SELECTION OF MOST REVIEW PROMISING DESIGN • CONFIGURATION REFINEMENT CONCEPT • DESIGN TRADE STUDIES NASA APPROVAL OF • DESIGN SENSITIVITIES CONCEPT SELECTION • PROPULSION EVALUATION • CONCLUSIONS AND RECOMMENDATIONS Figure i. - Work plan, hypersonic cruise aircraft propulsion integration study.

Therewasno special period of time allocated for generation of vehicle

configurations. Newideas for aircraft configurations, or for modifications

of existing concepts, were consideredthroughoutthe study.

3.1.3 Configuration screening. - All ideas for airplane designs were

considered and evaluated. There were two levels of screening; the initial level was essentially qualitative, the final was more detailed and provided quantitative data with which selected candidate designs could be compared.

The initial screening process was itself divided into two parts. All suggested design ideas were evaluated on a cursory basis to determine if there was sufficient merit in the concept to warrant further analysis. Naturally, some concepts did not survive this step. All too often the attractive fea- ture which led to the suggested configuration was obtained at the expense of penalties incurred in other features of the design. Where it was obvious the tradeoff would be unfavorable the concept was discarded.

There was also a comparison of designs, one with another. Those design concepts which appeared most favorable on the basis of this qualitative com- parison were laid out as three-view drawings in order to more vigorously assess their individual merit. In all, five candidate designs were treated in this manner. The design exercise permitted an evaluation of the practicability of the configuration, or permitted insight into the potential for making the design practical.

Such features as adequacy of room and safety for passenger accommodations, feasibility of integrating the two separate propulsion systems, potential for achieving a reasonably efficient structural design_ and the possibility of maintaining the proper relationship between center of gravity and aerodynamic center of pressure throughout the flight regime as required for vehicle sta- bility and control could all be assessed. In addition the aircraft was sized to a first approximation so that adequate fuel tankage was provided, landing gear could be located and its length determined to provide necessary tail scrape clearance, and the landing gear stowage problem conceptually resolved.

The design evaluation of the five candidate configurations led to selec- tion of two for final screening. One of these was the HT4 vehicle shape, pre- viously studied by NASA in wind tunnel tests. This shape was selected for two reasons; one, it appeared to be a very promising configuration (if certain modifications are made) and two, the existence of the wind tunnel data offered opportunity for verification of analytical results.

3.1.4 Vehicle synthesis. - The main tool used in the final screening and the trade studies is Lockheed's (Advanced System Synthesis Evaluation Tech- nique (ASSET)) program. ASSET is a vehicle synthesis model designed to size, parametrically weight, evaluate the performance, and cost large numbers of aircraft design options. A schematic presentation of the primary input and output data involved in the ASSET synthesis cycle, which is programmed on a high speed digital computer, is shown on figure 2. The ASSET program output consists of a group weight statement, vehicle geometry description, mission ii i w,_- _._ / lm • o ".

CANDIDATE CONCEPTS VEHICL _ITHESlS PRGGRAM (ASSET) PRELIMINARY DESIGN STRUCTURES WEIGHT SUBSYSTEM COST* AERODYNAMICS PROPULSION AND MATERIALS AND SIZING

[

MISSION PROFILE I VARY THE ASSET PROGRAM IT/W), (W/S) ETC PERFORMANCE REQUIREMENTS I PROGRAM OUTPUT

'-I PAYLOAD REQUIREMENTS I

I

COST" PERFORMANCE I SIZE WEIGHT I

I I I I

• RDT&E • BODY • MANEUVERABILITY • GROSS • INVESTMENT • WING • ACCELERATION • EMPTY PRODUCTION • TAI L • RATE OF CLIMB • STRUCTURAL TOOLING • OPTIMUM CRUISE • CRUISE ENGINES • MATERIALS SPARES & SSE ALTITUDE DISTRIBUTION DATA ETC • ACCELERATOR ENGINES • FIELD LENGTH TO • PROPULSION • OPERATIONAL • GEOMETRY AND LANDING MAINTENANCE • SUBSYSTEMS • FUEL CAPACITY REPLENISH SPARES • MAX GTO WEIGHT FUEL & OIL • OFF DESIGN PAY & ALLOW MISSION CAPABILITY • TOTAL SYSTEM COST *OPTIONAL Figure 2. - ASSET vehlc1# _ynthesls program schematic.

summary profile, and a summary of the vehicle's performance evaluation.

ASSET is composed of three major subprograms: vehicle sizing, performance evaluation, and costing (if desired).

k

Although the current ASSET program is very flexible and capable of analyz- ing many different types of aircraft, it was decided that significant changes should be made to more conveniently handle hypersonic aircraft because of the many propulsion forces involved and their interaction with the aerodynamic forces. A further complication is the change of these forces with angle of attack so that an iteritive solution is required for each point in the mission profile.

Accordingly, a new routine was written, to be used as a supplement to the existing ASSET program, which is called Hypersonic ASSET. This work was funded / t as a part of Lockheed's Independent Research and Dev, lopment (IRAD) program.

3.1.5 Candidate configurations. - From the matrix of conceptual designs suggested by both Langley and Lockheed personnel, five configurations were generated as candidates. These consisted of blended wlng-bodies, semi-blended wing-bodies and wlng-body. Both high and low wing were considered as well as various locat_on3 and arrangements of the baseline fixed geometry dual mode cruise propulsion system. These propulsion concepts have two things in common however; the use of a retracting inlet for the turbojet accelerator engine and the reduction of base drag by using a common nozzle for both the turbojet and scramjet exhaust. The retracting turbojet inlet is a major problem area in that it must have variable geometry when extended but retract into a minimum of space. Location of this inlet is also critical in that it should not inter- f_re with the scramjet during dual mode operation and should not be in an adverse flow region in particular at low speed and high angles of attack.

The general arrangement of the various HYCAT configurations are shown in the following figures: • Figure 3 HYCAT-I General Arrangement • Figure 4 HYCAT-2 General Arrangement • Figure 5 HYCAT-2 Cabin Arrangement • Figure 6 HYCAT-2 Cabin Cross Section • Figure 7 HYCAT-3 General Arrangement • Figure 8 HYCAT-4 General Arrangement • Figure 9 HYCAT-4 Propulsion Installation • Figure i0 HYCAT-5 General Arrangement ,-4 !

° I

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Wing V. Tail Area- Ft2 0644 (total) 1028 AR 1.357 .995 6O ALE DE6 15 30 ATE OEG Span- Ft 114.41 31.28 153.33 50.61 C R 15.26 13.68 C T MAC 103.14 35.68 T/C 3% 20 Wedge ___ ..... I........

LH2 tanks

L

Crew 45 Aux. compt.

/L (HT4) 0 .0375 .0667!.0937 .1832 .30G Ft. 0 11.4 20.4 28.6 FOLDOUT FPA;.:--_: \ i /'" I

1 i_ _/''"

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--.....

[ tanks . o , ....

Scramjet modules 227 235 I !._ Turbole _ Eqmp

Aux. compt. 114 I-" = _' _' I-" i

I_, 150,2201b H2; _', _ ; P • t I_.

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i L_

.403 .5495 .800 , .706 .800 .300 100 123.0 1G7.7 183.1 260 215.4 244.1

.'.: i' _ :._:.-._:J .L.n.OdlO"_ L..._

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------ ) ! -- _ . : _>¢

.11 1-1-I

_Z-..6---- t___.,_.....=-------- _"_ HT4 (REF.)

_ _j..__---- ..,...__..------'-'-" _ HT4 (REF.)

• ¢ I

\

,, / ":'-'

I

_ _ _x_- HT4 (REF.} / 227 235 /./ 300.5 !

[ i__ Turbolet Access 298 3 / "" / Wg = 646.130 T/W = .581 W/S = .67 (Ar,/SREFIT,I TM .0125 .- } ,'_==--z-- ......... --'_a_, _i ACTj = 99.9 " " ' -I": '_, ; ', l, i'TO.2301bH2 _;;' _ i __,--- -_"_2- _ ...... 7- - -- __.._ "f..----..... __' . ,, -'--L_- -- 2-D TJ I - " _-- Scramiet _--. Nozzle [_.

.800 .823 .9075 956 , 300 1.00 1 n.a_7 X/L.

334 Ft 39.5.4 244.1 272.5 291.T 305.1 Figure 3. - HYCAT-I general arrangement.

I OLDOU} .......... :_:,,

b

c Wing Tail Area- Ft2 3644 (total) AR 1.357 .995 ;\LE o ATE ° Span 144.41 33.23 153.33 52.6 CR 15.26 Cr .............. 14.22 t/c 3 20Wedge_ MAC 103.14 37.08 Bifurca /- - nonin '=:::_--_tank

ORIG!NAL P,_.GE IS

OF F3_:" ;'J':_'_'-'--'_

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iSta. 224.____

Biturcated("Pillow'_ _ • integraltank +36! _ " - Crew Fus. Ref.' _ .......... _ _ "'-_--- .... ---- Sta O(ft) .0375 0667 .0937 47.0 .1832 300 .403 HT4 _ Ref.)X_ L -_"_ Q' 11.4 ° zU.4 28.6 55.9 31.5 123.0 I i.//I/../" "_

/. /' \

i t

t __ BL 20.82 - .50 MAC \` S / ,\ // T]W = .58 '\ , _ 232.47 / AC/SRE F = .0125 Aox/A© = 2.9 _- J'V Turboiet (4)_ _ ' / ifurcated _._"----_ A c = _H/2 eng _ , ; .

nonintegral , _ ._/..._ /" ,n_ "_-'_-;-,. _ _ _2oTJ..

_ \ "_'_ t _-- f-"_l Nozzle ;_ _. __ __-_. _7- _. ___i thrOat . Fus. Ref.

_ .... - .... _L____ Scramjet- / _- Wing c_rrv-thrm _' _"'_--_J 0' 'Sta 244.6 'Sta. 286.2 iSta. 303.0 'nozzle

s,2. /

.Fusmge oreak -- _ Retracting T.J. inlet I ; / / ......41,. -_-- , _ "- .... ___L. _" : -_ _-_. :'_-- _'._, APU ----_/_ • . ..... _--_--__--=_, : .. _ _ .- :,,'_. - _ -__,._ _/_ T ,._L./___ ___ ' ___ :::., -_-__ _:__=____/ .... -4-- _ : ' I 4 - _ [ 177.0' .286.2 303.0 ,32.0-200.0 _ 224.0 244.6 320.3 334.0 .600 .706 .800 .893 .986 1.00 .5495 183.1 215-4 244.1 272.5 231.7 305.1 167.7 0 5 10 15 20 30 40 5U Scale 1" = 10' Figure 4. - HYCAT-2 general arrang_

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......... _--_:: •_--_-- 1-30,,".cARGO .I _ EQU,. L_'L__L_;_-

-_--.. L-_<LZ-- A/ran,- = 40 tt = 42 in. U _: "-'+"-'_-_=_ ...................

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/ -- 10 in.. 352 In, .... I (29.33 ft) l 0 10 20 3040 50 60 STA 192.0 STA 200.0 80 100 l _ , | . , i 1/60 Scale inches Figure 6. - HYCAT-2 cabin cross sectlon A qualitative comparison of the five configurations is shown in table i.

with the advantages and deficiencies of each listed for each criteria shown.

Each configuration has certain advantages but on balance the -i configuration was selected as the baseline reference because of the tunnel background data available as a check for our internal prediction methods. The -4 configura u tlon was selected as the first alternate configuration because of its favor- able propulsion installation, good low speed lift characteristics and the structural advantages of nearly circular fuel tanks and direct wing carry- thru structure. Disadvantages are the higher drag and weight of the exposed ..... prOpblsion installation, a higher wing weight and the added weight and drag of the horizontal tail.

3.1.6 Evaluation of selected candidates. - The two selected candidates (HYCAT-I _na -4) _ere optimized by means of the parametric data generated by the hypersonic ASSET program described in 3.1.4. The optimization procedure and resulting data are described in detail in Volume II. The propulsion systems used in both aircraft consisted of turbojets with retracting variable- geometry inlets, and fixed-geometry, dual combustion mode scramjet engines.

Other trade studies reported in Vol II consist of the effect of gross thrust deflection during cruise and the penalty incurred if the scramJet is not used in the Mach .9 to 3.5 region.

A weight comparison of the final optimized revision of both aircraft is shown in table 2. This table shows that the -4 conflguratlon requires a 42--percent increase in gross weight over the -I to accomplish the mission.

The reasons for this large difference are described in detail in Vol II, Sect. 4.3.

k, BOth of the configurations studied in this initial effort have certain advantages and deficiencies. These are magnified by the extreme growth sensitivity of the hypersonic aircraft to changes in inert or fuel weight.

.Table 3 lists the problem areas of each vehicle and suggested courses of action. While it is apparent that high drag and weight are bad, the modifica- tion of each configuration to exploit its best features is not so straight- forward. In fact, it may be that the melding of the best features of both configurr=ions may result in something similar to the HYCAT-2 configuration but with a means of obtaining a higher C L at low speed, in particular during landing.

3.2 Configuration Refinement The major conclusions drawn from this initial analysis of candidate configurations HYCAT-I and HYCAT-4 can be summarized as follows: --z ............ i ......

I' i'i I V. TAIL WING (PER SIDE) AREA - FT 2 9938.5 1400 ASPECT RATIO 1.50 1.0 ALE OEG 60 50 ATE OEG 35.78 26.15 SPAN 122.1 37.42 112.27 5O.52 CR CT 50.52 24.3 85.30 38.24 MAC T/C % 3 3.49 0.45 0.481 i

l

f f X/LREF 0 0.10 0.20 STA fit) 0 33.4 66.8

_ .-/ <

/ / i . /, / /- ,- J - / .... I "" _r .... " \ i- --1" ,, - __-_,__ _, , ' "r '" ...... -J-._-_ _ _T.-____---._, .... :':i:

ORIGINAL PAGE II "" \

U _ Intelral tank

_- __ _ ___-.__ _

, _ ..... i - -- -" .... _'- '---':_-_ - - .__ ' " "_ Equip

_--.__ _ /_,,_ .llHydll . \ 4- -. ' -,- -, '-j'

_-- -Ls;eF--?_ '=,J _*_ _

--- - _ax. coT;p_en, .53. -_..- Fu,.Joint /

• " ' -;\. I' Upper deck ," / T .... J _ _ -L-_ W -- ___ i Car _--_ - /"_ _- Scramjet _ "_ _--_//-- - Modules 0.30 0.40 0.50 0.68 0.70 0.80 100.2 133.6 167.0 188 200.4 205 ZR 233.8 267.2 GG _ etract Turbjet Inlet, FOLDOUT FRAME _ Range (Ac/Eng =22.6ft.2) / ,, / ./" / J / i / "'// , ' " _ Mach 0-3.5 . - _"_\ "_'-- Math 3.5-6 // \ WTO/SREF = 651b/ft2 FSLS (TJ)/WTO = 0.58 --__-'.-- \ __hrusl[ ;'Everse ACsj/SRE F = 0.125 _ Scramjetgeometric tegral tankr exit area= 310 ft2 Scram jets . / / " (AC/side= 62 ftz) _ View A-A -Heat_AP-tT_ . _ i Fus. Ref.

exch's_|-:.'_ --i [ i) I .... _-- _ ..j, Figure 7. HYCAT-3 general I 0 10 20 30 40 50 arrangement.

303.2 q'"' ...... J ' ' ' ' ' Scale- ft 0.80 0.90 s.00 (Ref.) 1/100 287.2 300.0 334.0 H tail i"" Win9 V. tail (exposed) Area - ft2 10 000 1495 600 Aspectratio 2.154 1.0 1.0 (perside) 60 50 60 ALE. - deg Span 146.75 38.66 Cr 120.6 55.24 28.87 Ct 15.68 22.09 5.77 Mac 81.6 .41._4_. 20,66.

Uc % 3 40 wedge 4o wedge .13 .4 .2 i .i i Sta. 60 St;, Integraltank 75560 _-_--+', ,_{, rus. rer.j / t___ u,¢w l_ ....... _ .......

I ' i i

I I i i f (t) i 1 1 ' I i Sta.-ft I 01 15 30 45 (_0 _'_' 1

FOLDOUT FRAME I

LOCKHEED

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LK 2f

-7

.50 MAC 146.75' I \\ W.t.O.

Sre f = 6

FSL__._3_S

WTO -.

ACSJ = .!

Sref 330 _4fl Sta-ft i 290 310 " :'-> ' _ ! 260 165 180 200 \_ 230 70 80 30 40 50 60 0 10 20 ft2 _- Scramiet(Ac = 63 i , ..__,l | r • I f.. , ..., Scale-ft.

Figure 8. HYCAT-4 general arrangement.

FOL/1_UT F_._,ME _

o /j "/ I °° i" 3.24 / ./J A =30ft2/eng(M =3.5) .J O // " i-

--------'Z-----"----

4 0 I ,mu oc= 2.6 M _= 3.5 oC=5-6 ° M<X=6

_r__OOUT _:RAMg

;OL_,..,:'_T..F. P, Ak_r;,

Pi_,CEDl.i_i PA'_-,Ei_,.._K NOT FILME_

6.6 0 !

5.4' AC = 122 (Tot)

/

...

.{ ...............

-----...._ s,!

I -" 7_,1 " o'<'_ ) _._..._.. _ _. _._..

I .+ + -- ............ i doorsand throat rev. exit --. Primary nozz' /------- M3.5 - 6 (SJ) Secondarynozzle and thrust rev.

/

1.00 C -b- , ._--_ M3.5 (TJ+SJ) T.O to M.5 1/4 0 0 5 10 Scale - Figure 9. HYCAT-4 propulsion InsCallaClon.

Jf / Wing V. Tail Area- ft 2 14 356 2050 AR 1.322 1.0 Retractable, Hi-lift, ALE - neg 80 - 65 60 low speed canard Span- ft 137.7E 44.72 269.16 63.89 C R 0 (theo.; 25.56 C T MAC 165.09 47.46 _, 0 (theo) 0.4 2.5 at fus 3.fJ at break 20 wedge t/c - % Integral Tank --,,, Crew (95 000 Ibs LH 2) "" Pas_angers Upper dec_ Fus. ref.

.u _Lo_v_er de_c, +10___t.lnd" ---- -Car_o STA 0 20 40 6O 8O 100 t_n 140 Sta- ft

PRECEDING ""-

• P=_-" BI..AhX NOT FILMED

FOLDouT FRAMF. I _, r 10 A/B C 36" pitch / _-Non i°te,,a, tank _j__..

, _ "-' , _,i:-',_ (FsLs ---- -- Win- stru_ _ _ J "t" Y c' Retractable / , / • , _'_ Sta 234 _ _ _ inlet ' ,we;_;:---:-_ - ............. - ........ . -___z_c_ . A ..... , . ] - ------

o _

140 160 180 200 300 340 349.5 0 10 20 30 40 50 60 70 80 100 --_' ------L--_._ _ L_____ CL 1725-.=; Hvoersonic configuration Fig" FOLDOUT F RA L-'E p

H

I

--Turbojets (FsL S = 93 000 each) Wg = 646 000 Z ACsj = 155 ft2 W/SREF = 45 Ib/ft 2 inlet----w / / SREF = 14 356 ft2 67' FSL----_S = 0.58 Wg

20...... JC__ Scr_je, ;6"

| 3OO 340 349.5 50 60 70 80 _. Ino J__ L____ '.5-.K Hvoerzonicconfiguration Figure I0.

HYCAT-5 general arrangement.

-DOUT -'.;.!;',_. -_

/

..._' • CON FIG U RATION Description Propulsion Integration HYCAT-1 Basic-FLT4 shape with = Favorable pressJre field bottom mounted pro- e goundarylayer growth pulsion. Tandem TJ = medium and SJ inlets. Forward e Unfavorable blockage of single deck pass, SJ inlet by TJ inlet compartment.

(M 0 to 3.5) >- e Aft underfuselage in jgt wake e Access to TJ's causes loss in volume.

Z Modified HT4 shape with HYCAT-2 e Favoraole pressure field for aft mounted propulsion.

LLJ SJ's ..J Top TJ inlets with SJ on QO = Tcp location of TJ inlet bottom. Mid-fuselage, will cau._eproblems due to ........

double deck pass.

boundary layer ingestion and compartment. Wing lower separation at low speed.

to permit struct-carry e Boundary layer growth = max.

thru.

e Good accessto TJ's.

= Remotelocation of thrust from C.G. accentuates trim problem.

High wing - Aft side Symmetric nozzle negates thrust mounted SJ.'s with TJ's vector trim problems.

on aft bottom.

e TJ inlets in favorable press ,,¢_, >.

Symlr, etric nozzle. Mid- field fuselage double deck pass.

e Wing-fuselage corner flow ipto compartment. Twin vert.

SJ not desirable.

tails.

e Weak press-field to SJ's- wing shock intersects inlet e Good accessto TJ'S.

HYCAT-4 Low wing - Wing mounted = Medium strength shock field.

propulsion with TJ's e Minimum boundary layer growth over wing - SJ's under e No TJ/_J inlet interference Area ruled fuselage with = Channel flow between fuselage

IPAX I '

double deck pass. com- and pods undesirable --L'Z partment. Conventional = Possible engine out tr!m vert. and horiz, tail.

problem (supersonic) " --- tr e Close coupling of thrust and C.G.

e Good accessto T,_'s and SJ's.

Low wing double delta.

e Same comments as for HYCAT-2 Aft mounted propulsion above.

HYCAT j_ y

with top TJ inlets - SJ's ,,¢_,>.

¢._ ¢:Z on bottom. Sears-Haack IPAXl _ semi-blended body with double deck pass. com- partment between tanks Canard for low speed trim.

FOLDOUT FRAME I

PI_CEDI.NG P,_.: =,_.,.',,_ t_OT

Aerodynamic Characteristics Structural and Tankage Volumetric Efficiency Passenger Location _r e Lift of flattened forebody e Limited wing box carry • Loss in volume due to e Good access for Ioad;qg ?

contributes to high hyper- TJ access thru - load taken by and serving ( sonic LID.

frames or integral tanks • Single deck max.

e Not protacted by wing • Difficult to get C.G. for e Gear must retract into compartm_ntcadses structure 3000 ft 3 vol. loss enough forward to nlatch aero wing.fairing required.

= Max. C. G. travel center.

= Fuel tank weight penalty compared to double deck for pillow tanks e Lateral directional stability adequate • Lowspeed C L limited - no high lift- low AR.

• Same asabove -1 except that • Same as above -1 except Better than -1 above d_Je e Over-wingaccess required wing must be moved aft to wing has direct carry to double deck pax.

= Protected by wingstructure counteract required shaft of thru.

compartment. Min, C.G. travel C.G. with aft propulsion. Some volume loss in e Fuselage deepened to permit propulsion area.

double deck max. compartment (higher drag) e Fwd tanks circular - o For•body wave drag high due to Large loss in volume due = Good access SJ inlet flow field contraction minimum wt.

to gear stowage e Not protected by wing desired.

e No wing carry thrn - structure • Tip fins may have undesirable weight penalty e Vulnerable to gear collapse interaction at low speed. = High wing requires e Min C.G. travel long, heavy gear.

e Added drag due to exposed I Direct wing carry • Very good-(Prop, not in e Over-wingaccess required nacelles and horizontal tail thru - min. wt.

fuselage) • Partial protection by wing • Wing bending relief due • Small volume Ioss due to structure e Good Iowspeed C L due to flaps and drooped ailerons to propulsion location e Min. C.G. travel gear stowage • Horiz-tail causes e Minimum trim drag - long tail arm e Horizontal allows use of flaps fuselage bending loads e No particular aero advantage • Direct wing carry thru - Good - some loss in pro- e Moderate over-wing accessrequired min. wt.

unless inboard panel L.E. couZd pulsion area • Partial protection by wing be made subsonic e Circular fwd. tanksand structure pass. compt. - min. wt. e Uin. C.G. travel e Lower C L makes airport performance critical • Gear retracts into wing • Canard required for rotation - fairing required and trim at low speed.

EOLDOUT -F,_AM _ TABLE i. CANDIDATE CONFIGURATION COMPARISON Producibility Index* Comments Passenger Location Selected as baseline reference I 1.0 e Good accessfor loading (Baseline Value) because of tunnel data and and serving • Not protected by wing previous studies structure Tandem inlet not acceptable ........

e Max. C. G. travel • revision required 0.8 e Over-wingaccess required = Expected to be s;mi!ar in • Protected by wing structure performance to -1 • Min. C.G. travel = TJ inlet location marginal 0.9 • Good access • Scramjet inlet location marginal e Not protected by wing structure • Vulnerable to gear collapse e Min C.G. travel 0.7 • Selected as 1st alternate e Over-wingaccess required e Partial protection by wing configuration e Good low speed characteristics structure e Min. C.G. travel may negate lower cruise LID • Body lift could be increased by chines or flattening of body 0.75 • Moderate over-wing access required e Potential of hypersonic double delta not known e Partial protection by wing structure • Could evolve to hypersonic e Min. C.G. travel arrow wing?

*Structure only - no equipment (lower value = lowest mfg. cost) TABLE 2. - COMPARISON OF OPTIMIZED IPICAT-I AND -4 -1 -4 GeneralCharacteristics: kg/M2 (Ib/ft 2) Wing loading 373.5 (76.5) 488.2 (100) Thrust/weight daN]kg (-) 0.49 (0.50) 0.44 (0.45) 0.011 (0.011) CapturearM/wingarea 0.012 (0.012) Weights kg (Ib) Grossweight 307 382 (677 649) 435 196 (959 426) Total fuel 108453 (239094) 164 140 (361 860) Fuelfraction .3528 .3528 0.3772 .3772 Payload 19 051 H2000) 19 051 (42 000) OEW 179 877 (396 555) 252 005 (555 565) Std plusoperatingitems 6 611 (14575) 7 065 (17 560) 173 265 Empty weight (381 978) 244 O4O (538 006) Stfucture - fraction .2517 .2517 0.2893 .2893 Wing 22402 (49 387) 48920 (107649) Tail 2 631 (5 800) 3 783 (8339) 36282 Body (79 987) 48005 (105 831) 11 716

Ldg.gear (25829)

15551 (34283) Surfacecontrols 2720 (5 997) 3662 (8073) 1 622 Nacalleandeng.section (3 576) 5979 (13 180) Propulsion fraction .2106 .2106 _1943 .1943 Engines (T.J.) 20 276 (44 701) 25 837 (56960) Air induction(T.J.)

4 145 (9 138) 5313 (11 713) Scram jets 9724 (21 437) 11 489 (25329) Fueltankage and systems 30 127 (66410) 41 323 (91 100) 0.1014 Systems, furnishings and equip.- fraction 0.1014 0.0772 0.0772 MissionPerformance: 5.21 CruiseL/D (average) 5.21 4.72 4.72 Cruise specificrange km]kg (n.miJ]b) 0.1425 (,0349) 0.0878 (.0215) Descentrange km (n.mi.) 891 (481) 600 (324) 52 267 Blockfuel required kg (Ib) (203 410) 143 302 (315 921) FAR T.O. fld. dist. m (ft) 3 016 (9 895) 2 118 (6 950) FAR Ldg.rid. dist m (ft) 3 203 (I0 510) 3 182 (10 440) 5 971 (10 494) Energyutilization kj Btu 9 274 (16 298) seatkm (seatn.mi.)

P_CF./,)?_ p,_,_ _'_-_*'_K NOT FILlCVE D

TABLE 3. - CONFIGURATION PROBLEM AREAS POSSIBLE PROBLEM COURSES OF ACTION RESULT HYCAT-I: 1. Turbojet inlet blocks S.J. e Move inlet to top aft of fuselage e Marginal region f_r TJ inlet operation Mach 0.8 to 3.5 o Modify config.

2.

Passenger compartment Move to mid- fuselage and double e _Bettprv(!l. _ff ciency . .

vol. not efficient.- cg deck passengers e Reduce cg travel travel too large z Add canard 3.

e Added weight and reduced wing lift Low value of CL e Use elevons asflaps z cg must be controlled or a horizontal durin9 low speed T.O, and tail is required Ldg.

HYCAT--4: Decrease AR I. Heavy wing wt.

e Decreases low speed C L Decrease leading edge sweep a Reduces wt. and high speed drag 2. Propulsion drag and weight Bury TJ's in fuselage- put inlet on e Oecr.o.ase fus. volume top or bottom e ;ncreases TJ base drag to the Mach 3.5 - 6.0 region e Moves cOaft o Add chines or flatten fuselage 3. Increase fuselage lift • Noncircular fuel tanks (added wright) • Lessefficient fuel volume The landing field length is the critical sizing constraint.

Turbojet accelerator engines should be buried within the airframe when they are not used. This serves to minimize both drag and nacelle weight.

Q The arrangement of the propulsion system in I_CAT-I blocks the scram- jet inlet in the Mach 0-3.5 flight regime. The inlet retraction and stowage concept Ks too complex.

Lift provided by a.flattened vehicle forebody (or by use of strakes) is important to improve hypersonic L/D.

Wing weight is critical in that higher aspect ratios, while providing higher low-speed lift, incur an excessive weight penalty.

The use of a horizontal tail (or canard) is required to provide trim for relative changes in center of gravity and aerodynamic center. A further advantage is that it allows the use of drooped ailerons (flaperons) for low speed lift.

The forward passenger compartment location on HYCAT-I is not efficient and the center of gravity movement is too large.

Consideration of the above conclusions in the initial effort resulted in the selection of the basic HYCAT-I shape for modification and refinement because of its aerodynamic efficiency at cruise. The following modifications were made: • A new propulsion configuration was generated to overcome the objec- tions of the HYCAT-I arrangement.

• The passenger cabin was moved to mid-fuselage in a double-deck arrangement similar to that shown for HYCAT-2.

I A horizontal tail and wing flaps were added. This alleviates, to some extent, the low speed lift disadvantages of a low aspect ratio wing.

The final baseline configuration designated HYCAT-IA is shown in fig- ure Ii. This is the starting point for the design trade studies reported in detail in Volume II, and is the configuration on which the propulsion studies described in the following section were conducted.

3,3 Propulsion Concepts As the primary focus of the study, two propulsion concepts were evaluated: i.) a concept with a variable-geometry inlet and turbojet engine and a separate fixed-geometry inlet and scramjet engine and 2.)

a :oncept with a variable-geometry inlet supplying air to both a turbojet and a ramjet engine. The supersonic combustion cycle was used with the fixed-geometry inlet since the scramJet cycle is less dependent on variable geometry to achieve the proper inlet throat area over the required speed range.

3.3.1 Separate Inlet, Turbojet-scramJet syst_, - This concept is shown schematically in figure 12. It consists of a variaEle-geometry, retractable inlet for the turbojet engine and a fixed-geometry inlet, combuster and nozzle for the scramjet. The dual mode engine uses thermal choking by means of heat addition in the subsonic combustion mode from Math .9 to Mach 4.5. Super- sonic combustion is initiated at Mach 4.5 and is continued to Math 6 for use throughout cruise. The turbojet is used for landing, takeoff and accelera- tion to the scramjet takeover point at Math 3.5 to 4, at which time the turbojet inlet is retracted as shown. A common exit nozzle is used for _oth the turbojet and scramjet. Advantages of this concept are: • A simple fixed-geometry cruise engine with no moving parts reduces complexity.

The supersonic mode reduces the engine heat load and internal pressure due to reduced static temperature and pressure in the inlet, combustor and nozzle.

r

• Potential for operation at higher Mach numbers such as Mach i0.

The disadvantages are: The exposed fixed-geometry scramjet causes large installation drag in the critical transonic region as well as during subsonic ....

cruising flight (cold flow drag).

The fixed-geometry of the scramJet limits the inlet air flow capa- bility at lower Mach numbers.

The turbojet inlet retraction requirement causes problems in mechani- zation and sealing.

3.3.2 Common variable-geometry inlet, turbojet-ramjet system. - This system is shown schematically in figure 13. The method of operation is similar to that of the turboJet-scramjet combination with the exception that a common inlet supplies both the turbojet and ramjet up to Mach 3.5 at which time the turbojet is shut off and only the ramjet is used up to and including cruise. The interior surface of the inlet aft of the cowl, the ramlet dif- fuser, and the ramjet module are all regeneratively cooled by the hydrogen fuel.

The advantages of the ramjet compared to the scramjet are: • Lower installation drag at low supersonic and subsonic speeds.

• Inlet retraction is not required as it is for the turbojet inlet of the turbojet-scramjet system.

• Higher thrust in the supersonic and low hypersonic speed regime due to the variable inlet and nozzle.

• Less development risk and facilities requirements.

The disadvantages are: • Higher unit heat flux at cruise due to the subsonic mode of operation (near stagnation pressure and temperature).

Limited in maximum flight Mach number. A rapid deterioration in thrust and impulse occur at speeds higher than the Mach 6 of this study, compared to the scramjet.

A further comparison of the two systems will be found in latter sections of this report.

I i

Fwd tanks I & 2: 97.98 Sta 147 90 700 Ibs LH_ -_ L,

I

Ref i 2.99 4fz

1___2

I -- .Sta ft 0 11.49 20.43 28.71 56.13 31.31 123.46 .7832 .J.30 .0375 .0667 .0937 X/L (Ref. HT4) O .403

FOLDOUT FRA_.,F

n

[] =_

[] 0

i i

t '

12_o"L_ .'_- \ I /

J°°

-17 .07 13 .1 .

I I 6 I I

i

seat pitch

i

Aft tanks 3 & 4: a 147 Sta 168.34 Sta 216.29 i

I

1.16 330 Ibs LH 2 ,,_j_1,. = ; .................. I Ip ,_.. - - , -_ I'; Lowerdeck] ] L, _,........ (, 1 -" '.:_'_; Cargo -";-- "_ L: :' -_"_ :; ..... --_'-'_ "

......... ... _ _,,._. _ .,__---:_---i --_

123.46 168.34 183.82 216.29 \ \ 245.09 . 273.58t278.02 292.88 306.36 .403 ...... :54-95 ...... :l_l_0 -- .71}1_ .......... X _-:800 ...... __-:3'-.9075 _956 1.00 Turbojets (4).-_ _---- Scramjet Modules (10}' -75 0001b sis (uninstid) ___ = i _ . . _-- 0 '= 20 30 40 FOLDOUT Fp,/l.t.,,. _; f- ._.

Gross wt (est) 600 000 Ib J

'NinqIoadinQ 85 Ibs/ft2 Thrust/wt (sis) 0.50 Fuelwt. 207 030 Ib ,ft tanks3 & 4: ;_278.02 292.88 306.36 • 9075 ..956 1.00 1.10 Scale: 1" = 10' L.... i I I I ..__._..:.i__

0 '° 20 30 40 _o 6b--'---_----_---_

,J 80 90 100 Figure ii. General arrangement baseline version of HYCAT- IA.

._OL _eVOL;T • ="_"'";" / 35 /////_F Fuselage countour Fuselage boundarylayer _o* _o* _, / l- Rampangle

- , / __ _.., _.._w __.___tl J_j --_ J._ _ _

PRECEDING .PACE E:,./_'E{, r_OT ._

_" _J''_ '_"_ _.... i " _" -' "_,'_._i __ (2 REF Plane p. "i,J,o.o°.

A2SJ(varies) -- _',

\'q

!

Turbojet "on," inlet upen- MOO0-3.5 ...... r , Turbojet "off" inlet closedMOO3-5-6 Figure 12. Separate inlet, turbojet- scramJet system scnematic.

Turbojet I / / /_ Retractuuleboundary / / loayer flap .,-.--..- / ,-- 60* (M 3.5) Fu_ contour /! .

FOLDOUT _RAME

Jf -- Turbojet shutoff doors- closeat roach3.5

\

\

-l_.l_l Vii" ..... _L

_,........... ................................... t ......... i

_'__.a__________ Ac = 8.36 m2 190 tt 21(diverter closed) Ac/module = 2.09 m2 (22.5 ft 2)

\

FSL S (TJ| = 333 600 N (75 000 Ib) (uninstld.} \ X Ramlet A3Rd = 3.72 m2 (40 ft 2) A 6 = 16.72ra2 (180 ft 2) ft 3- 4 5 1]} 0 I 2 Scale _J_ I J ' i,i I IX = ' J m Figure 13.

C_mmon Variable - Reometry inlet, turbojet-ramjet system schematic.

r

Options in location of the turbojets in relation to the cruise engines and location on the aircraft were examined in the configuration definition phase and were shown in Sect. 3.1.2. The final location of the turbojets adjacent to the cruise engines was dictated, however by the necessity of using a common nozzle for both in order to reduce the base drag of an unfill d nozzle in the critical transonic and low supersonic speed regime. The loca- tion on the aircraft was a result of aircraft c.g. requirements and the rotation (scrape angle) required during takeoff and landing.

A detailed description of the installation and performance of both propulsion concepts is presented in Section 3.3 of Volume II.

4. BASIC TECHNOLOGY 4.1 Aerodynamics Volume II, section 3.1 contains a detailed discussion of the methods, analysis and data on the aerodynamic characteristics.and stability of HYCAT-I, -4 and the final revision of HYCAT-IA.

4.2 Aircraft Weight Estimation Volume II, section 3.4 describes the methods and assumptions used in the airframe weight prediction. Propulsion weights are discussed in Sec- tion 3.3 of Volume II.

4.3 l_,itial Propulsion Data The turbojet-scramjet propulsion system was used in the initial screen- ing phase. A detailed discussion of the basis for selection of the turbojet and scramjet engines, data sources and installed performance can be found in Volume II, section 3.2 for this phase of the study.

PRECEDING PAGE _tANK NOT FILMED

4.4 Final Propulsion Evaluation In the final phase the turbojet-scramjet system configuration was revised and performance was recalculated. The major changes were as follows: The turbojet inlet and scramjet were located on a ramp to allow con- current operation of both in the Math 1.0 to 3.5 region. This also allowed more nozzle area and minimized the volume loss in the fuselage.

Flow field viscous effects on mass flow were included in the scramjet performance after the turbojet boundary layer diverter was closed at turbojet shutdown.

The inlet contraction and mass flow ratio schedule was revised to account for the increased external contraction and decreased local Mach number resulting from the ramp.

The installation and performance of the alternate propulsion concept consisting of turbojets with separate modular, subsonic combustion ramjets, both using a common inlet, was provided.

The vehicle flow field, inlet characteristics, installation losses and installed performance of both propulsion systems are described in section 3.3 of Volume II. The weight estimates for both concepts and estimated cooling requirements for the ramjet system are also included in the same section.

5. COMPARATIVE ANALYSIS OF PROPULSION SYSTEMS As in the initial phase, the hypersonic ASSET program was used in a systematic optimization of the variables of wing loading (W/S), thrust-to- weight ratio (T/W), and capture area to wing size ratio (Ac/S) in all trade- off studies. The criterion for selection was minimum gross weight and the major constraint was the 10,500 ft maximum takeoff or landing field. FAR international fuel reserve requirements were used except that 5% of the fuel used at the end of cruise was used in lieu of 10%. No limitation was placed on airport noise in this study.

5.1 Separate Inlet TurboJet-ScramJet System The turbojet-scram jet final optimized point design aircraft selected to perform tile Math 6, 200 passenger, 9260 km (5000 _l.mi.) mission is shown in figure 14. In summary, the essential features of this final version compared to the ttYCAT-1 of Phase I are: • Incorporation of a horizontal tall for stability.

• Revision of the propulsion configuration as decribed in section 3.3 of Volume II.

• Incorporation of the passenger compartment in a double deck, arrange- mcnt in the center fuselage.

Table 4 summarizes the geometry, weight and performance characteristics.

A listing of selected ASSET program printout pages can be found in Appendix A.

Table 5 is a summary of the unit structural weights based on total planform for wings and call and wetted area for the fuselage. The thermal protection system weight shown is an average weight. Some of the windward qurface_ will require higher weights and leeward less than shown. The thermal protection system could be either an active or a passive type.

5.1.1 Weight sensltivity. - An investigation was made of the selected point design HYCAT-IA to changes in systems, propulsion, or structural weight items. This would occur during final design if for example, the wing weight were to increase 2000 Ibs. If the aircraft were to perform the design mission carrying the same payload it would have to be resized. The resulting change in gross weight would be 5.27 kg of gross weight per kg of original weight change, l.e; a "growth factor" of 5.27. Thus the original wing weight [n_rease of 907.2 kg (2000 Ib) wouhl cause a gross weight increase of 3656 kg _IO 3So lb:_ which wolild involve all non flxed-weight items.

5.1.2 Fuel sensitivity. - The sensitlvltv to changes in the total fuel load was also investigated. This could be caused, for example, by a degrada- tion dr, ring design of propulsion efficiency or a change in reserve fuel requirements. The analysis, using ASSET to resize the aircrafc, showed that an orL_ina[ iucrease of I kg of ft,el required woi,ld cause a 6 kg increase in the gro.qs w_ight, it is not st, rprising that this sensltiv[ty or growth fa_tor TABLE 4.

- NYCAT-IA POINT DESIGN CHARACTERISTICS Turbojet-Scram jetSystem - 200 pa_engefs - Mach 6 92.60 km (5000 n.mi.) Range t/c = 3% ALE = 65 o AR = 1.357 GEOMETRY: m 2 (ft 2) 816.8 (8792.1) Wing Ref. Area m 2 (ft 2) 383.8 Wing Exposed Area (4131.5) m (ft) (388) Fus. Length m (ft) (24.46) Fus. Equivalent Dia.

m 2 ((12) 887.1 Fus. Planform Area (9 549) m 2 (ft 2) Fus. Wetted Area 2 402.1 (25 857) m 2 f2 (t) 11.03 Scramjet Capture Area (118.7) m 2 (ft 2) Horiz. Tail Total Area 177.8 m 2 (ft 2) 115.2 Horiz. Tail Exposed Area m 2 (ft 2) Vertical Tail Area 90.2 (971) Kg (Io) WEIGHTS: Gross Wt. 350 953 (773 706 Fuel: Block 107 038 (235 975 Reserve 19 085 (42 074 Total 126 123 (278 049 19 051 Payload (42 000 7 050 Oper. and Std. Items (15 542 198 729 Empty Weight (438 116 Structure: 106 026 (233 744 24 276 Wing (53 960 Tail 6 857 (15 117 • 41 337 Body (91 131 13 023 (28 711 Ldg. Gear Surf. Controls 3 046 (6 716 Thermal Protection 15 407 (33 966 1 852 Nac. and Eng. Sect. (4 083 75 286 (165 974 Propulsion Engines (Turbojets) 23 133 (51 037 4 948 Air Inlet (Turbojets) (10 909 3 281 Fu,_land Oil System (7 234 31 531 (69 512 LH 2 Tanks, Insul. and Supports 530 (1 169) Eng. L;ontrols and Starter 11 845 Scran'jets (26 113) 17416 Furn., _q.i_. and Subsystems (38 396) PERFORMANCE kg/m 2 (Ib/ft 2) 429.6 (88) Wing loading 0.49 SLS Thrust/Weight daN/.g 0.50 0.0135 Capture/Wing area 0.0135 Far takeoff ft dist. (Eng. (t) m (ft) 2 568 8 426 Cruise L/O (average) 5.17 5.17 (Ib/hr/Ib) Cruise SFC (average) kg/hr/daN 1.43 Cruise alt. m (It) 29-30175 95-99,000 m (ft 3 225 Far landing dist. 10 580 (keas) m/s 95.7 Approach speed 186 kJ/seat km (Stu/seat n.rni) 6 927 Energy consumption 12 174 4_ HORIZORTAL WING VERTICAL EXPOSEO CHARACTERISTICS BASIC EXPOSED BASIC 123511 97109 -- AREA t_ FT) 87821 41814 [ 1913.S, 1.0 ASPECT RATIO 1.957 2.0116 31.10 SPAN IFTJ 109Z3 62.04 49.03 ROOT CHORO _FT.) 140 4.t 13.39 132g TIP ('HORO IFT J 14._0 12 29 TAPER RAl'10 (_1 2488 34.58 MAC .FT) 26.5.1 34 56 6O 9klEEP (0EGJ _ T_ 4.0 T/C ROOT C_ 3.0 4 0 a,0 TIC TIP 1%) 3 0 ,..

LH2 i 8 + ,__ 14, Fwd Tanks 120 987 Ibs LH2 . Section A

\

FOLDOUT -FRAk_ 1 t

s.c, on _._,o.B ° ee_ A"Tan_ ,ST0_0_sLH2_ 7/

_c

_-- . _J___. _T 0 ;_o.._[- ..... - FuelWeight 278 047 'bs -"_. ,=O _:.,.._jf'- -'- /'__i ]'hrustjWeight (SLS)..50 389.36 f't " " Figure 14. HYCAT-IA, final general arrangement.

IF.I_LDOU T -FPAI/F

TABLE 5. - UNITSTRUCTURAL WEIGHTS, HYCAT-IA

Wing Kg/m 2 (Ibs/ft2) 29.97 (6.14) Horizontaltail 21.97 (4.50) Verticaltail 32.71 (6.76) Fuselage (includingLH2 tanks) 30.37 (6.22) Thermal protection* (average) 14.11 (2.89) *Basedon exposed planformareasof wing,tail andfuselage is higher than the weight growth factor (above) since it also involves an increase in the fuselage weight to carry the fuel. To further illustrate the above effect, if the propulsion system SFC were anticipated to degrade by 2 percent in service, a not unreasonable assumption, the original gross weight would have to be increased by 15 150 kg (33 400 Ib) or the payload decreased by approximately 3130 kg (6900 Ib) if the gross weight were not increased and the same range held.

5.1.3 Range sensitivity. - Using the ASSET program to resize the aircraft the original design range of 9260 km (5000 n.mi.) was reduced to 8334 km (4500 n.mi.) and 7408 km (4000 n.mi.), holding the prime constraint of land- ing field distance constant. The primary effect of course is the reduction in fuel fraction with the secondary one being the decrease in wing loading required to meet the landing distance as the block fuel fraction decreases with range. Table 6 lists some of the characteristics of the aircraft de- signed for each range. The table shows a growth sensitivity of 136.4 pounds of gross weight per nautical mile between 4000 and 4500 with the sensitivity increasing to 177 between 4500 and 500 nautical miles.

5.1.4 Subsonic cruise range. - If the 9260 km (5000 n.ml.) point design HYCAT-IA (Wg - 350 953 kg (773 706 Ib)) were to cruise at subsonic speeds with a full fuel load and the same reserve fuel requirement the maximum range would be 6267 km (3384 n.mi.). The optimum Mach number is 0.90 and the cruise altitude is from 7920 to 8534m (26 000 to 28 000 _t.). _ The average cruise L/D is 8.31 with an SFC of 0.498 ib/hr/Ib which gives us an average range factor (M(L/D)/SFC) of 15 compared to 21.7 for the Mach 6 cruise case. This is not surprising since the subsonic L/D of such an aircraft would not be expected to be high (12-15). Turbojet engine used in the study is also not the best engine for subsonic operation. If an SFC of 0.34, _lich would be equal to that of turbofan engine could be obtained, the range would approach 9260 km (5000 n.mi.). This of course suggests the dual cycle engine being studied for application in the SCAR program. The range could also be improved by reduction of the propulsion drag in this region.

PRECEDI_ P_.GE BLANK NOT FI_ o TABLE 6a. - POINT DESIGN CHARACTERISTICS OF HYCAT-IA AT RANGES OF 7408, 8334 AND 9260 km (S.I UNITS) T/W = 49 daN/kg,--_ = .0135, t/c - 3%, AR = 1.357 Range-km 8334 9260 GrossWeight kg 279 915 310840 350 953 Fuel: Block 77 893 90 688 107 038 Reserve 15 348 16 974 19 085 Total 93 241 107 662 126 123 Payload 19051 19 051 19 051 Oper. and Std. Items 6 240 6 595 7 050 Empty Weight 152 312 177 531 178 729 Structure 85 690 94420 106 026 Propulsion 59 099 66 157 75 286 Furn. Equip.and Systems 16 594 16 954 17 385 Fus. Length m 103.33 _ 110.03 118.26 Wing Loading kg/m2 CruiseL/D 4.97 5.06 5.17 CruiseAir. m 29.3-30 180 28.96-30 180 28.96-30 180 Far T.O. Dist m 2 509 2 568 Far Ldg. Dist. m 3 158 3 179 3 225 ApproachSpeed m/g 94.4 95.0 95.7 Block Time hr 1.95 2.08 2.21 EnergyConsumption kJ/seatkm 6 301 6 521 6 927 GrowthSensitivity kg (Wg) km 33.41 43.33 5.2 Common Variable Geometry Inlet, Turbojet-Ramjet Systems The approach used in the design optimization of the turbojet-ramjet propulsion system consisted of replacing the turbojet-scramjet system with the weight and performance characteristics of the turbojet-ramjet system using the selected point design scramjet aircraft described above 350 953 kg (773 706 ib). This was done to obtain a "side by side" comparison of the weights and fuel consumption for each system in the same configuration. Both aircraft have a thrust-to-weight of 0.50 and a wing loading of 429.6 kg/m2 (88 ib/ft2). The only difference is that while the scramjet had an optimized

TABLE 6b. - POINT DESIGN CHA_'_CTERISTICS OF h_'CAT-IA AT RANGES

OF 4000, 4500, AAYD 5000 N.MI. (CUSTO._L_RY UNITS) T/W = .50,,_ = .0135, tic = 3%, _ = 1.357 Range - n.mi.

5000 (REF.)

Gloss Weight -Ib 617 097 685 273 773 706 Fuel: Block Reserve 171 720 199 930 235 975 37 421 Total 33 837 42 074 237 351 Payload 205 557 278 049 42 000 42 000 Oper. andStd. Items 42 000 13 757 14 540 15 542 EmptyWeight 355 784 391 383 Structure 438 116 Propulsion 188 912 208 158 233 744 Furn. Equip. and Systems 130 289 I45 848 165 974 36 583 37 377 38 326 Fu=. Length ft 339 Wing Loading Ib/ft2 82 CruiseLID - 4.97 5.06 5.17 CruiseAir. ft 96-99 000 95-99 000 95-99 000 Far takeoff dist. f_ 8 017 8 230 8 426 Far landingdist. ft 10 362 20 430 10 580 Approach Speed keas 183.5 184.7 Block Time hr 1.95 2.08 2.21 EnergyConsumption Btu/seatnm 11 074 l 1461 12 174 GrowthSensitivity Ib (Wg) 136.4 176.9 n.ml.

capture-to-wing area ratio of 0.0135 (Ac = ii 03 m 2 (118 " =_-- - system Ac/S ratio selected • - . o • _ .i _r-)), the • l, as 0 01_75 based on obtazning the same netr_h3ruest ts[ztl t down.

as the scramjet at turbojet The ASSET progrmn was not allowed to size the aircraft but simply flew the airplane through the mission holding the takeoff gross weight co,lstant.

A summaz-v of the weight output is shown in table 7. Inspection of the table shows that the equipment, structural, standard, and operating weight items are almost identical, but that the propulsion systenl is 466 kg (10293 ib) lighter. This is due primarily to tilat fact that the turbojet- scramjet requirL, s a separate inlet for the turbojet.

TABLE 7a. - WEIGHT BUILD-UP COMPARISON OFTURBOJET-RAMJET SYSTEM INSTALLED

IN TURBOJET-SCRAbtJET POINT DESIGN AIRCRAFT (S.I UNITS)

T/W = .49 daN/kg W/S = 118.26 kg/m2 Ac/S = .0135 (SJ). 0.01275 (RJ) _W T J-RJ (SJ-RJ) TJ-SJ 17 427 -10 17417 kg Equipment -508 (106 534) (106 026) Structure 24 476 24 476 Wing 6 403 6 734 Tail 42 075 41 364 Body 13 023 13 023 Ldg.Gear 3 046 Surf. Controls 1 852 1 773 Nac.& Eng. Sect.

15 407 15 407 Thermal Protec.

(70 616} 4669 (75 285) Propulsion 22 163 23 150 Engin_ 4 948 9 407 Inlet 32 545 31 531 FuelTanks 3 282 2 304 FuelandOil System 53O Eng. Contr.& Starter 1 690 11844 Scram_,zs/Ramjet 194 678 198 729 EmptyWeight 7 145 -95 7 050 Std. & Oper.Items 19 051 19051 Payload 130 179 4057 126 122 FuelWL Available 350 953 350 953 GrossWei_zt A weight advantage of 987 kg (2176 ib) is also shown for the Mach 3.5 turbojet used with the ramjet vs tile Mach 4.0 turbojet required with the scramjet system. The significant end result of the weight build-up is that tile ramjet system has an advantage of being able to carry a fuel load 4057 kg (8944 ib) more than the scramjet system. Note that the higher body, lower tail, and higher tank weights of the ramjet system are due to the longer body required to contain this extra fuel weight.

A comparison of the mission fuel consumption is shown in table 8. The right hand column shows that the advantage in fuel consumption is 6804 kg (15 000 Ibs) of block and 2223 kg (4900 ib) of reserve fuel for the ramjet system. The difference during climb and descent is mainly due to the lower transonic propulsion installation drag and higher specific impulse of the ramjet system (See Section 5.3). Tile descent fuel flow of the scramjet could be decreased at the expense of the descent range due to the higher propulsion TABLE 7b. - WEIGHT BUILD-UP COHPARISON OF TURBOJET-RAMJET SYSTEM INSTALLED IN TURBOJET-SCRAHJET POINT DESIGN AIRCRAFT (CUSTOS[ARY UNITS) T/W = .50 W/S = 88 Ac/S = .0135 (SJ), 0.01275 (R J) ,_W TJ-SJ TJ.RJ (SJ-R J) Equipment (38 398) (38 4_3) "22 Structure (233 743) (234 864) "1121 Wing 53 960 53 960 Tail 14 117 14 845 Body 91 190 92 757 Ldg. Gear 28 711 2J 711 Surf. Controls 6 716 Nac.& Eng. Sect.

4 083 3 909 Thermal Protec.

33 966 33 966 Propulsion (165 973) (155 680) 10 293 Engines 51 038 48 860 Inlet 10 909 20 738 Fuel Tanks 69 512 71 748 Fuel and Oil System 7 235 7 284 Eng. Conlr. & Starter 1 169 1 120 Scram jets/Ramjet 26 t!?

5 930 l:.mpty Weight 438 114 428 964 SId. &Oper. Items 15 542 15 762 -210 Payload 42 000 42 OOO Fuel Wt. Available 278 047 286 991 -8 944 Gross Weight 773 706 773 706 J drag which would result as explained in Section 4.2.1.6. Again the reserve fuel advantage is due to the lower propulsion drag of the ramjet system during l, the subsonic cruise. During the cruise portion of the mission something of an anomaly occurs in that while the specific impulse of the ramjet is 3008 see., that or: the scramjet is only 2518 sec. (16.3% lower); however, the specific r,n£e of the scramiet vehicle is only 2.1 percent lower. A small part of this is due to the higher average gross weight (1.97%) of the ramjet aircraft in cruise but the major difference is in the propulsion-aero force account- ing. As was pointed out in Section 4.2.1, the turbojet inlet and the scram- jet are mounted on a ramp to allow concurrent operation of both. Thus the "propulsion system" includes this ramp even when the turbojet inlet is closed.

Except for the turbojet inlet, the ramp forces would have normally been in- cluded in the aerodynamic forces but are all charged to propulsion resulting in the apparent low specific impulse of the scramjet. The ramp forces in cruise are included in the spillage drag and lift as well as the smaller spillage drag and lift forces of the scramjet inlet itself.

[

!

b

TABLE 8a. - COMPARISON OF MISSION FUEL CONSUMPTION OF TURBOJET-RAMJET SYSTEM INSTALLED IN POINT DESIGN TURBOJET-SCRAM JET AIRCRAFT (S.I UNITS) T/W = .49daN/kg W/S = 118.20Ng/m 2 Ac/S = 0.0135(SJ),0.01275(RJ} _WFuE L TJ-SJ TJ-RJ (SJ-RJ) GrossWt. (Takeoff) kg 350 953 350 953 Takeoff & Climb to M6.0: Opt. CruiseAir m 29 - 302.00 27.4 - 28.400 Fuel Used kg 48 737 44749 3 988 Dist.

km 2 008 2 069 Cruise LID 5.17 5.20 ISP* daN/kg/sec 2 469 m 2 950 Fuel Used u,,.. kg 54 324 52840 1 484 Dist.

km 6 091 6 052 m km]kg .1121 .1145 Descent.." (M6.0 to 128.6 m/g) Fuel Used kg 2 938 1 638 1 300 Dist.

km 1 101 I 141 Air Maneuver& Ldg: Fuel Used kg 1 039 987 52 Total Block Fuel 107 038 100 214 6 824 ContingencyFuel: 5% of Block Fuel kg 5 352 5 011 341 Climb Fuel Used kg 3 214 2 962 252 Dist. km 5O Cruise Fuel Used kg 4 705 3 388 1 317 Dist. km 5,.., Descent Fuel Used kg 105 -11 Dist. km 30 Min. Loiter & Ldg.

Fuel Used kg 5 722 5 418 304 Total Res. Fuel kg 19 086 16 883 2 203 Total MissionFuel kg 126 124 117 097 9 027 *Defined asnet thrust in flight axisdirectiondividedby total fuel hold .f

Im •

| TABLE 8b.

- COMPARISON OF MISSION FUEL CONSUMPTION OF TURBOJET-RAMJET SYSTEM INSTALLED IN POINT DESIGN TURBOJET-SCRAM JET AIRCRAFT (CUSTOMARY IJ_ITS) T/W = 0.5 W/S = 88 Ac/S = 0.0135 (SJ). 0.01275 (RJ) y _WFUEL TJ-SJ TJ-RJ (SJ-RJ) GrossWt. (Takeoff) - Ib 773 706 773 706 Takeoff & Climb to M6.0: Opt. CruiseAir ft 95-99 000 9_93 000 Fuel Used -Ib 107444 98653 8 791 Dist. n.mi.

I O84 1 117 Cruise L/D 5.17 5.20 ISP* sec.

2 518 3 008 Fuel Used Ib 119 763 116490 3 273 Dist. n.mi.

3 289 3 268 NM/lb 0.02746 0.02805 Descent: (M6.0 to 250 kts): Fuel Used Dist.

Air Maneuver& Ldg: Fuel Used 2 290 2 176 Total Block Fuel 235 975 220930 15 045 ContingencyFuel: 5% of Block Fuel Ib 11 799 11046 _ Climb Fuel Used Ib 7 085 6 529 Dist.

n.mL Cruise Fuel Used Ib 10 373 2 903 Dist.

n.mL Descent m rJ Fuel Used Ib -24 Dist.

]] n.mi.

30 Min. Loiter & Ldg.

Fuel Used Ib 12 614 11945 Total Res.Fuel Ib 42 076 37 220 4 856 Total Mission Fuel Ib 278 051 258 150 19 901 J *Defined asnet thrust in flight axisdirectiondividedby total fuelflow i I The final result is that while the scramjet is charged with a higher spillage drag it also provides a very high spillage lift contribution to the aircraft. This is shown in the table 9 comparison of the ramjet and scramjet baseline aircraft, each cruising at its optimum altitude. The final result is a slight advantage of 4 percent in specific range for the ramjet system which is partially negated by the 1.97 percent higher average cruise weight of the ramjet aircraft.

TABLE 9a. - BASELINE AIRCRAFT CRUISE COMPARISON OF THE TURBOJET-RAMJET AND THE TURBOJET-SCRAM JET SYSTEM (S.I. UNITS) LW -- .49_g W/S = f18.26kg/m 2 Wg = 350953 kg ACsJ 11.03 m 2 = 10.41 m 2 = ACRJ TJ-SJ TJ-RJ O Cruise wt.

kg 286 339 285 792 _uise Air.

m 29 261 27 737 C_uL_eL]D 5.18 5.2 Angle of attack rad 0.0745 0.0675 O Centrifugal lift -kg 14 889 14 760 Propulsion: JB(Gross Thrust deflection) rad 0.0873 0.0873 Capture area m 2 11.03 10.41 Gross Thrust -kg 187 772 178951 Inlet drag 54O 2 747 Momentum drag 134 411 125 162 Spillage Drag 3 030 S¢illage lift 41 287 Total Propulsion Uft 52 012 18 837 Q AereL Lift Req'd. = O " O

®

219 529 252 188 Aero. Drag- Q /L/D 42 381 40 456 Net Thrust in Fit Axis 42 381 48 456 Fuel Flow kg/sec 16.833 10.107 dan Isp Fuel Flow 2 469 2 950 Spedfic Range- km/kg 0.1074 0.1118 TABI_E 9b. - BASELINE AIRCRAFT CRUISE COMPARISON OF THE TUR:_O.IET-RA._ET AND THE TURBOJET-SCR._IJET SYSTEM (CUST,Y.IARY I'N]TS) T/W = 0.5 W/S = 88 Wg = 773 706 Ib ACs J = 118.7 ft 2 ACR J = 112.1ft 2 TJ-SJ "J'J-RJ O CruiseWt. -Ibs 631 258 630 053 Cruise NL ft 96 000 91 000 CruiseLiD 5.18 5.2 Angleof attack deg 4.27 3.87 Centrifugal lift -Ib 32 823 32 540 Propulsion: (Gross Thrustdeflection) -deg Capturearea ft2 118.7 11ZI GrossThrust -Ib 413 960 39_ 750 Inlet drag I 190 Momentumdrag 296 320 275 930 SpillageDrag 6 680 Spilhge ;ift 91 020 Total Propulsion Lift 114 664 41 528 @ Aero. Lift Req'd. = O O " @ 483 970 555 970 Aero. Drag- Q /L/D 93432 106 825 Net Thrustin Fit. Axis 93432 106 825 Fuel Flow Ib sec 37.11 35.51 I = Net Thrust sp Fuel Flow sec 2 518 3 008 SpecificRange- n.mi./lb 0.0263 0.02739 o Following a checkout of the performance and weight of the baseline air- craft described above, the synthesis program was allowed to size the turbojet- ramjet aircraft to provide the @esign range capability of 9260 km (5000 n.ml.)

for a matrix of various thrust-to-weights, capture areas, and wing loadings.

The minimum gross weight aircraft that meets the landing field distance con- straint was then selected. A summary of this point design is shown in table i0.

As anticipated from the lower propulsion weight and fuel consumption of the ramjet system, the gross weight shows a 72 576 kg (160 000 Ib) reduction compared to the scramjet system. A lower wing loading was required to meet the landing field length constraint because of the reduced block fuel fractions.

5.3 Comparison of Separate Inlet and Common Inlet Systems The cause of the difference in the point design gross weights of the optimized scramjet and ramjet systems can best be shown as in table II expressed in terms of weight fractions. As can be seen items such as payload, operating items, furnishings and subsystems tend to remain constant in weight and as a result, increase in weight fraction as gross weight decreases. The structural fraction remains almost constant with the major change being in the propulsion -- r.

and fuel weight fractions which decrease by 1.86 and 1.06 percent of gross weight respectively for the ramjet system. This is a total reduction of 2.92 percent and using the weight sensitivities given in Sections 5.5.1 and _° 5.5.2 one could have predicted that the final gross weight would be in the 272-283 500 kg (600-625 000 Ib) range.

The most significant actual causes for this weight decrease are the reduced propulsion weight and fuel consumption of the ramjet system. As already stated, the low speed propulsion installation drag of the scramjet system is the most important single factor. This is shown by a comparison of the mission climb history shown in figure 15 for the turbojet-scramjet systems compared to figure 16 for the turboJet-scramjet at the same gross ° weight. The thrust-drag pinch points occur in the Mach 1-1.5 region and at the end of turbojet operation. The higher installation dlag of the scramjet °-- in the transonic region is shown as is the lower thrust at the end of turbojet operation (Mach 4 to 5). It should be explained that the initial intent was L to terminate turbojet operations at Mach 3.5 but it was found that a deficiency in the thrust available from the scramjet occurred at the end of turbojet operation. Two alternatives were considered: I) increasing the capture area by approximately 20% or 2); extending the turbojet operation to Mach 4. The first solution is undesirable because of the weight penalty of 2268 kg (5,000 ib) involved. The second alternative was selected and t!le perfor- mance envelope of the turbojet extended to Mach _ by assuming that the turbo- jet airflow would be reduced at Mach 4 so that the turbojet inlet would not TABLE I0. - HYCAT-IA POINT DESIGN CIlARACTERISTICS Turholet-Ramjet System • Mach 6 • 9260 km (5000 n.mi) range t/c : 3% ALE = 65° ,oR : 1.357 Geometry: m 2 (ft 2) Wing Ref. Area 662.4 (7 129.9) Wing Exposed Area m 2 (ft 2) 280.4 (3 018,2) Fus. Length m fit) 105.1 (344.9} Fos. Equivalent Oi•. m fit) 7.46 (24.46) Fus. Planform Are• m 2 (ft 2) 735.7 (7 919) Fus. Wetted Are• m 2 (ft 2) 2 043.5 (21 997) Inlet Capture Area m 2 (ft 2) 8.12 (97.4) Horiz. Tail Total Area m 2 (ft 2} 146.8 (I 580) Horiz. Tail Exposed Are• m 2 (ft 2) 90.49 (974) Vertical Tail Area m 2 (ft 2) 74.49 (801.0) Gross Wt

kg 0b)

278 136 (613 174) Fuel: Block 83 778 (184 696) Reserve 13 236 (29 179) Total 97014 (213 875) Payload 19051 (42 000) Oper. and Std. Items 6 328 • __L _- (13951) Empty Weight 155 743 (343 349) Structure: (84 682) (186 688) Wing 18 199 (40 121) Tail 5 661 (12 481) Body 33 870 (74 670) Ldg. Gear !0 839 (23 895) Surf. Controls 2 497 (5 505) Thermal Protection 12210 (26 918) Nac, and Eng. Sect.

1 405 (3 098) Propulsion (54 476) (120 098) Engines (Turbojets) 17 428 (38 722) Air Inlet 7 339 (16 179) Fuel and Oil System 2 831 (6 242) LH 2 Tanks and Insul, and Supports 24 254 (53 469) Eog, Controls and Star_er (887) Ramjets 2 096 (4 620) Furn,. Equip and Subsystems 16 585 (36 563) TABLE I0. - HYCAT-IA POINT DESIGN CHARACTERISTICS (toni'd) Perfo.-mance Wing Loading kg/m2 (Ib/ft 2) 419.9 (86) SLS Thrust_Neight daN/k9 - 0.49 (0.50) Capture]Wing Area - - 0.01225 (0.01225) Far T.O. Dist. (Eng. Out) m (ft) 2 557 (8 390) CruiseL/D (Average) - - 4.93 (4.93) CruiseSFC (Average) _v/daN (Jh]lr]lb) 1.218 (1.194) Cruise AlL " m (ft) !7.4-28 650 (90-94 000) Far LandingDisL m (ft) 3 172 (10 406) ApproachSpeed. m]s (keas) 94.9 (184.5) EnergyConsumption kj (Btu/) 5 422 (9'529) seatkm have to be larger than at Mach 3.5. Because of the higher operating pressure and temperature however, the weights of the inlet and turbojet were increased 4.56 and 4.46% respectively. The final specific weight of the turbojet inle_ including boundary layer and retraction mechanism is 595.6 kg/m 2 (122 ibs/ft ) and the sea level _minstalled static thrust-to-weight of the turbojet is 7.58 (assumed constant with size).

The final result is an increase in the climb fuel required for the scramjet system of 48 737 kg (107 444 ib) compared to 44 740 kg (98 633 ib) for the turbojet system. In order to isolate this effect, the scramjet air- craft was resized by making the assumption that the total propulsion installa- tion drag of the scramjet system was exactly equal to that ot the ramjet sys- tem. The results of this assumption are shown in the third column of table Ii which indicates a dramatic weight reduction of almost 36298 kg (80 000 ib).

5.4 Turbojet-Ramjet System With Fixed Diverter The previous analysis of the turbojet-ramjet system assumed that the variable-geometry inlet and ramjet combustor could function while ingesting the fuselage boundary layer in the Mach 3.5 to 6 region _diverter closed).

Since this assumption cannot be established short of test validation, an analysis was made to determine the effect on propulsion characteristics and aircraft weight of a fixed dlverter. The diverter was a vee-shaped ramp de- signed to plow off the maximum boundary layer displacement thickness. The effect of the diverter was to increase the inlet recovery by decreasing the viscous losses in total pressure and to increase the mass flow by removing the displacement thickness. The disadvantages are an increase in drag a_d °.

f

TABLE ii. - WEIGHT FRACTION COMPARISON - POINT DESIGN, SCRA_tlET AND RAMJET ;YSTEMS Low Drag TJ - SJ TJ - RJ TJ -S J* 350 953 278 136 315 204 GroseWeight(Ref.) - kg (613 174) (695 070) (Ibs) (773 706) Fractions: 0.0543 0.0685 0.0604 Payload 0.0200 0.0228 0.0211 Std. and Oper. Items 0.O496 0.0596 0.0539 Fum., Equip.andSystems (0.1239) (0.1509) SUBTOTAL (0.1354) 0.3021 0.3045 0.3030 Structure 0.2145 0.1959 • 0.2129 Propulsion Fuel: 0.1390 0.1504 0.1336 Takeoff andclimb 0.1548 0.1437 0.1574 Cruise 0.0113 0.0071 0.0119 Descent and landing (0.3051) (0.3012) 0.3029 Total Block 0.0544 0.0476 Reserve 0.0458 (0.3504) (0.3488) Total Fuel 0.3487 0.5739 0.5447 0.5615 Propulsion plusFuel Fraction *With propulsioninstallationdragbelow Mach2 equalto turbojet-ramjetsystem.

weight. The installed performances compared to the retracte_ azverter zs shown in Figures 17, 18, and 19 which indicate an increase in thrust but a decrease in specific impulse in the Mach 4 to 6 region both in full power and part power cruise at Mach 6 as shown In Figure 19.

A weight penalty was caused by the increase in total pressure recovery which increased the inlet weight by 11.3% and the ramjet module weight by 11.9%. A further penalty was caused by the fixed diverter, the surfaces of which were assumed to consist of a metallic heat shield over high temperature insulation. This penalty was partially offset by the removal of the retrace- able diverter panels and actuators. The final specific weight comparison is: Retractable Fixed Diverter Diverter kg/m 2 of Ac (ib/ft 2 of Ac) kg/m 2 of Ac (ib/ft 2 of Ac) 903.1 (185) 1045.7 (214.2) Inlet Specific Wt.

258.3 (52.9) 289.0 (59.2) Ramjet Specific Wt.

I161._ 237.9 1334.7 (273.4) Total Specific WT.

Net jet thrust(FNJ) fNet thrust (FN) 18 - 400 '

L_ J

16 --

1 '

J T/W = .50 14 - Ac = 118.7 ft.2 3OO 12 -- Z tO -_ 200 j drag 6 -- ( Jrcra t drag 4 -- I00

I

_/',J

2 -- T.J. on =_ ° I S.J. on .

0/

I I I

2 4 Freestream Machno. - Mp Figure 15. Mission climb history, turbojet-scramjet system.

6O ..< 3.5 3.4 3.3 __ e '-, 3.1 O.

Figure 16. Mission climb history, turbojet-ramjet system.

r •|

26 I I

m m No boundarylayerdiverter " With bouqdarylayerdiverter

2.4 [ /_-" ____ _s o_Angleoflaztack.deg

4 ° -, _, 1.8 _ .

1.2 _ .8 • 1000 qootrajectory .4 e Eq. ratio (_ = 1.0

2 ]

3.5 4.0 4.5 5.0 5.5 b.0 Machno.- Moo Figure 17. Installed thrust coefficient ramjet with fixed dlverter.

°- L 2 4.0 4.0 _ ,., No boundary layer diverter , With boundary layer diverter 3.8 : • 3.8 3, 3.4

//

3.2 / 3.2 8 o Angle of attack - deg Z 3.0 3.0 1000 q Trajectory Eq. ratio (_) = 1.0 2.8 2.8 2.6 3.5 4.0 4.5 5.0 5.5 6.0 Mach no. - M oo Figure 18. Installed specific impulse ramjet with fixed dlverter.

Angleof attack -deg z No boundarylayerdiverter " "" "" " "_ With boundarylayerdiverter 0 ='Max. power(_ = 1.0) Mach6.0 28 956 m (95 000 ft) 2.6 .6 .8 1.0 FN Net thrust coeff CFN = ----- q. AC Figure 19. Cruise part power performance - ramjet with fixed diverter.

Thefinal diverter propulsion characteristics andnewweight werein-

corporated into the ASSET vehicle synthesis program and the aircraft (HYCAT-IA) was reoptlmlzed. The results are listed in table 12 which shows that the increase in gross weight of approximately 3 percent is mostly due to the in- crease in propulsion weight with the decrease in fuel specific impulse being largely offset by the increased ramjet thrust available in the Mach 3.5 to ........... 9 region. In summation, it appears that should a dlverter be required for the turbojet-ramjet system that the penalty in terms of aircraft growth would not be excessive.

6. STUDY CONCLUSIONS In an aircraft that is operated in a conventional manner, i.e., takeoff to cruise to descent and landing, the off-deslgn characteristics are of equal importance to the cruise performance. This is particularly true in the hyper- sonic transport due to its high growth sensitivity to weight and fuel consump- tion. This is emphasized in this study when one compares the propulsion characteristics that contributed to the final difference in the gross weights of the fixed and variable geometry systems. The fundamental reasons for the difference are due primarily to the following: 6.1 Installation Drag Figure 20 shows a comparison of the individual drag items that make up the total installed propulsion drag. It is obvious that the major item is the spillage drag of the fixed geometry engine. The reason for the difference is that the variable geometry system with a common inlet can supply the air- flow demands of both the turbojet and rmmjet and in so doing reduces the spillage airflow to about 35 percent of the total as shown in Figure 21. In contrast, the fixed geometry system with separate inlets for both the turbo- jet and scramjet must spill about 65% of the total forebody streamtube which results in a much larger drag penalty which, in turn, requires a combination of more tu£bojets, or more capture area and/or higher fuel consumption during L acceleration. A further penalty is incurred during subsonic cruise (reserve requirement) due to the high cold flow drag of the scramjet.

t J TABLE 12a. - AIRCRAFT WEIGHT COMPARISON OF TURBOJET - RAMJET SYSTEM WITH RETRACTABLE AND FIXED DIVERTER (S.I. UNITS) Range= 9260 km Fixed Retractable Diverter Diverter Weights kg 278 136 286 448 Grosswt.

100211 97 014 Fuel 19 051 19 051 Payload 6 328 6410 Oper. and Std. Items 160 776 155 743 Empty Weight 86 911 84 682 Structure (57 185) (54 476) Propulsion: 18451 17 564 Engine(Turbojets) 7 339 Air Inlet 2 831 2 906 Fuel & OilSystem 25 053 24 254 LH2 Tanks. Insul. & Supports Eng. Controls & Starter 2 242 2 096 Ramjets 16 680 16 58E Fur.. Equip.and Subsystems Characteristics 424.7 419.9 W/S ks/m2 0.49 0.50 T/W daN/kg 0.0115 0.01225 AC/SREF m 2 8.120 Ac m2 0.3488 0.3498 Fuelwt. fraction 0.1996 0.1959 Prop.wt. fraction 0.5494 0.5447 Total fuel & prop. fraction i" r | TABLE 12b. - AIRCRAFT WEIGHT COMPARISON OF TURBOJET - RAMJET SYSTEM WITH RETRACTABLE AND FIXED DIVERTER (CUSTOMARY UNITS) Range= 5000 n.mi.

Retractable Fixed Oiverter Diverter (Ib) Grosswt 613 174 631 500 Fuel 213 175 220 924 Payload 42 000 42 000 Oper. andStd. Items 13 951 14 132 343 349 Empty Wekjht 354444 Structure 186 688 199 803 Propulsion: (120 098) (126 069) Engine(turbojets) 38 722 40 677 Air Inlet 16 179 17 880 6 242 Fuel& Oil System 6 407 53 469 55 232 LH2 Tanks, Insul& Suppeels Eng. Controls& Starter 887 Ramjets 4 620 36 563 Fur., Equip.and Subsystems 36 773 Characteristics WPS T/W O.50 0.51 0.01225 0.0115 Ac]SREF

(h2)

87.4 83.47 Ac Fuel wl. fraction 0.3488 0.3498 Prop. wt fraction 0.1959 0.1996 Total fuel& prop.fraction 0.5447 0.5494 5O ACsJ = 8.918 m2 (96.1 ft2) • = 8.361 m2 (90 ft2) - - ACTj _ S w • FSLSTJ = 133 440 dan (300 000 Ib) S_ ..¢ - / ,#B_ .,= 30 ' 20

s

w__,_ DSPILLAGE

.o....-- z

-- 0 ! _%,J I t DBASE

,oI , !

1o

0 .------ _ "--'-" 30 m .,_ "_-(LL 60 _" "" " ,, _ ,,,,., J 5O DBASE + DINLE T _ DINSTL = DSPL+ 4O i 2O

i ,°

J 0 1000:1oo Trajectory O,_-,,-,, TJ+RJ

- __ lo

O" " " TJ+SJ

I I

O-- 0 1.6 .6 .7 .3 .9 1".0 1.1 1.2 1.3 1.4 1.5 Mach no. - M oo Figure 20. Installation drag comparison TJ-RJ and TJ-SJ systems.

..... I"- 35% Spillage . _ Spillage 1.0 TJ + RJ spillage .8

_1_

_1 _ TJ + S,J spillage _' .5 .4

.3 i l i

2.5 .5 1.0 1.5 2.0 M_ Figure 21. Inlet mass flow comparison, TJ-RJ and TJ-SJ systems.

L_ _..y- 6.2 Thrust Available, biach 3.5 to 5 As described In section 5.3 a thrust deficiency in the scramjet system occurred at the end of turbojet shutdown. This required that the turbojet ._° operation be extended to _hch 4. A further penalty in climb fuel consump- tion-followed after turbojet shutdown in the Math 4 to 5 region. This is shown in figure 22 which illustrates the lower thrust and [sp of the scramjet compared to the ramjet system. This is directly attributable to the fact that while the mass flow capacity of both systems is approximately equal, the total capture area of tile turbojet plus the fixed geometry scramJet is 14.31 m 2 (154 ft 2) compared to 8.36 m 2 (90.6 ft 2) for the common inlet of the turbojet-ramjet systems as shown at the top of the figure. The lower mass flow ratio capability of the scramjet consequently causes an increase In spill- °.

age drag as indicated by the lower net Isp of the scramjet.

6.3 System Weight Comparison ° A slde-by-slde comparison of the propulsion systems weights was prepared by holding a constant gross weight of 317 520 kg (700 000 lh) and using the optimum thrust to weight and Ac/b values determined for the final point de- sign turbojet-scramjet and turbojet-ramjet (with fixed dlverter) systems.

Table 13 shows that while the sum of the common inlet plus the modules is only slightly more than the scramJets ii 475 kg (25 238 ib) compared to i0 839 kg (23 896 ib), the turbojet-scramJet r_quires a separate turbojet [n]et with a total net penalty of 43 39g kg (9 566 ib_ or 13.6 percent heavier than the ramjet system. Also shown in the table is the total fuel fraction plus tankage fraction for each system. The bottom llne shows that the total weight penalty for the scra_jet compared to the ramjet system is a gross weight fraction of .0257 or 8 160 kg (17 990 Ib) at a constant gross weight of 31 7520 kg (700 000 ibs). This difference in weight then, considering the - i_ .;- growth factor accounts for the final difference in gross weights of 350 953 kg (773 706 lbs) for the turboJet-scramJet and 286 448 kg (631 500 Ibs) for the turbojet-ramjet systems with fixed diverter.

In summary, the essential difference of the svstems, is not in the com- bustion mode (subsonic vs supersonic) but is due to: I.

The reduction in both mission fuel consumpti_in and installed propul- sion weight made possible by the use of a common varlable-geomety inlet for both the turbojet and ramjet engines. The reduction in spillage drag of the common inlet in the critical transonic region allows a smaller cowl size and reduced fuel consl,mption both in acceleration and subsonic cruise.

7O -,.--_.4-_-_,,._._'_ ¸ _ _ _._.--"_---" _ ._ _.° 1.0!

RJ Spillage

' 1

.8

"'-- j

SJSpillage .4 oJ .2 Scram jet Ramjet O, =4 0

_1

!

Z _ m _ND m ammzm, 2.5

'_ 8 1.5

_ Z .1_ u.

ii Z _ •,o Subsonic mode - ' " v!

.5 .0 3.0 3.5 4•0 4.5 5_0 Freestream Machno. _- Moo Figure 22. Performance comparison, Mach 3.5 to 5; TJ-RJ and TJ-SJ systems.

/ TABLE 13. - PROPULSION SYSTEHS WEIGHT COHPARISON "-" -./ Gross weight = 317 520 kg (700 000 Ib) ._ = 424.7 k_9=(71 Ib/ft 2) m Z SRE F " 747.5 m 2 (8046 ft 2) TJ-SJ System TJ-RJ SystemO Characteristics: daN T/W (optimum) Ac/S (optimum) -_g (') 0.49 (0.50) 0.50 (0.51) Ac REF - 0.0135 (0.0135) 0.0115 0.0115 m2 (ft2) 10.09 108.6 8.596 (92.5) Weights: daN Turbojet FSLsJWTtd "_" (') 7.433 (7.58) Turbojet wt.

_.767 (7.92) kg (Ib) 20945 __ (46 174)

20447

__ (__45 076) Scram jet system: TJ inlet specific wto

(Lb-__)

km _ ft2 595.6 (122) TJ capture area m2 (ft2) • 7.516 (80.9) TJ inlet vat ko (Ib) 4 477 (99870) SJ module specific wL kg Ib _ 1074 (220) SJ internal area m2 (ft2) 196.2 (2 112) SJ _ight ko (Ib) !0 839 _ _)---- Ramjet system: Common inlet specific wt 1 045.7 Inlet wt. (214.2) kg (Ib) 8 990 RJ module specific wt. O (19 820)

k (Ib)

650.3 (133.2) RJ internal area (ft 2) RJ wt 38.09 kg (Ib) (410) 2 485 (5 478) Total propulsion wt .-_ _ kg (Ib) 36 261 (79 940: Total propulsion un fraction 31 922 (70 374) 0.1142 (0.1142) Total fuel + tankage wt. fraction 0.1005 (-I 005) Sum of propulsion + fuel 0.4493 (0.4493) 0.4373 (0.4373) and tankage fractions 0.5635 (0.5635) --...-.-..

0.5378 (3.5378) (_/ith fixed diverter '_c: .444 or A 3 = (41.1 ft 2) 2.

The use of this variable£geometry inlet increases the inlet air flow (and thrust) in the critical Math 3.5 to 5 region after turbojet shutdown.

The net result is that the turbojet-scramJet system is penalized in both fuel consumption and Installed weight caused by high subsonlc/transonic spillage drag and by low thrust in the Math 3.5 to 5 region due to a lower mass flow resulting from the fixed geometry scramJet engine.

Other c0n¢l_sign _ reached in the configuration study phase are a_ follows: • The gross weight of aircraft to perform the design mission are in the 272 160 to 362 880 kg (600 000 to 806 000 Ib) class.

The lift provided by a flattened fuselage forebody is important in improving hypersonic L/D and in providing the flow field and geometric width necessary for the propulsion installation. This is of particular importance in hydrogen-fueld aircraft with a large potential fuselage to wing planforms area ratio.

The use of a horizontal tail in the selected configuration was re- quired for trim purposes and "paid its way" by allowing the use of drooped ailerons to obtain more low speed llft with the final payoff being the reduction of wing size and weight. A further benefit is the reduction of the neutral point variation with Math number.

The most critical design criterion is to meet the landing field length constraint without increasing the wing aspect ratio or reducing the wing loading, both of which options result in increased gross weights.

The propulsion system should be integrated with the fuselage to avoid excessive wave and friction drag. It should also be located far enough forward for balance purposes and to allow for takeoff rotation without requiring a long main gear for clearance. Further benefit is of the reduction of propulsion moments when the system is located near the center of gravity, and a reduction in the boundary layer displacement thickness. Adverse effects of the fuselage boundary layer could dictate the use of wlng-mounted propulsion nacelles.

The location and optimum inclination of the gross thrust vector can make a significant reduction in cruise fuel flow by reducing the aerodynamic llft required and subsequently the drag.

Based on supersonic transport design experience and the high growth sensitivity of the hypersonic transport, the imposition of airport noise constraints would have a very adverse impact on vehicle size although it is possible that this could be mitigated to some extent by a variable cycle accelerator engine in which, as secondary bc_nefit, the subsonic SFC could be improved thereby reducing the reserve fuel consump t Ion.

7. STUDY RECOMMENDATIONS The primary recommendation, considering the propulsion application to a transport mission, is to pursue the use of a common inlet for the acceleration and cruise engines and to provide a higher thrust level in the Mach 3 to 5 region by variable geometry or other means.

The majority of the remaining recommendations stem from uncertainties in the prediction methods used in the study. Testing andanalytical correlation is required in the following areas: • Demonstrate that either the variable or fixed _eometry engines (inlet + comhustor + nozzle) could operate efficiently while ingesting the boundary layer from the long fuselage forebody.

• If a dlverter is required for either system what is the low speed drag and what lift contribution is caused by the shock fleld imping- ment on the fuselage or wing underside?

• Determine by test the spillage llft and drag forces in the transonic region.

• Simulate propulsion flows co determine base drags and moments.

• Further analytical work is required to define the comparative weights and cooling requirements of both propulslon systems.

7_ REFERENCES io G.D. Brewer, R.E. Morris, G.W. Davis, E.F. Versaw, G.R. Cunnington, Jr., J.C. Riple, C.F. Baerst, G. Garmong, "Final Report - Study of Fuel Systems for LH 2 - Fueled Subsonic Transport Aircraft," NASA CR-i45369, Lockheed- California Company for NASA-Langley Research Center, July 1978.