/

NASA Technical Memorandum_ 1061 !8 .......AIAA-92-5080

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The Design and Evolution of the BetaTwo-Stage-to-Orbit Horizontal Takeoff

and Landing Launch System

Leo A. Burkardt

Lewis Research Center

Cleveland, Ohio

and

Rick B: Norris

Wrigh t L a bora toryWright-Patterson Air ForceBase, Ohio :r

Prepared for the ....................

Fourth!ntemational Aerospace Planes Conferencesponsored by the American Institute of Aeronautics and Astronautics _

Orlando, Florida, December !-4, 1992

:7 _

N/ A

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(NASA-TM-I06118) THE DESIGN ANDEVOLUTION OF THE BETA

TWO-STAGE-TO-ORBIT HORIZONTAL

TAKEOFF AND LANDING LAUNCH SYSTEM(NASA) 18 p

N93-27018 '

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https://ntrs.nasa.gov/search.jsp?R=19930017829 2020-05-17T03:45:04+00:00Z

THE DESIGN AND EVOLUTION OF THE BETA TWO-STAGE-TO-ORBIT HORIZONTAL

TAKEOFF AND LANDING LAUNCH SYSTEM

Leo Burkardt*

National Aeronautics and Space AdministrationLewis Research Center

Cleveland, Ohio 44135

and

Rick Norris

Wright Laboratory

Wright-Patterson Air Force Base, Ohio 45324

Abstract

The Beta launch system was originallyconceived in 1986 as a horizontal takeoff and ATF

landing, fully reusable, two-stage-to-orbit,manned launch vehicle to replace the Shuttle. It EWwas to be capable of delivering a 50,000 lb.payload to low polar orbit. The booster GLOWpropulsion system consisted of JP fueledturbojets and LH fueled ramjets mounted in HSCTpods in an over/under arrangement, and a singleLOX/LH fueled SSME rocket. The second HTOL

stage orbiter, which staged at Mach 8, was

powered by an SSME rocket. A major goal Ispwas to develop a vehicle design consistent withnear term technology. The vehicle design was .IPcompleted with a GLOW of approximately2,000,000 lbs. All design goals were met. L/DSince then, interest has shifted to the 10,000

lbs. to low polar orbit payload class. The LHoriginal Beta was down-sized to meet thispayload class. The GLOW of the down-sized LOXvehicle was approximately 1,000,000 lbs. Thebooster was converted to exclusively air- SSMEbreathing operation. Because the boosterdepends on conventional air-breathing SSTOpropulsion only, the staging Mach number wasreduced to 5.5. The orbiter remains an SSME TSTO

rocket-powered stage.

Nomenclature

Advanced Tactical Fighter

empty weight

gross lift off weight

I-IJgh Speed Civil Transport

horizontal -takeoff-and-landing

specific impulse {(lbf. sec)/lbm)

hydrocarbon jet engine fuel

lift to drag ratio

liquid hydrogen

liquid oxygen

Space Shuttle Main Engine

single-stage-to-orbit

two-stage-to-orbit

*Member AIAA.

propellantmassfraction{propellantmass/(GLOW- payloadmass)}

Introduction

Since the beginning of the "Space Age" andbefore, interest has been strong in air-breathingHTOL launch vehicles. The advantages seemobvious; the air provides the oxidizer,

significantly increasing I sp, and winged vehiclesusing runways remove the need for speciallaunch platforms, thus allowing simplifiedlaunch operations and more versatile basing.Both SSTO and TSTO vehicles have been

examined in the past. Unfortunately, thepropulsion and materials technology were notadvanced enough to design a practical vehicle.

While air-breathing HTOL SSTO vehicles arestill a formidable challenge, propulsion andmaterials technology have caught up with theTSTO making it an attractive concept. The Betalaunch system is such a TSTO concept. Thispaper will discuss the design guidelines andcharacteristics of the Beta launch system and itsevolution from Beta I to Beta III.

Design Goals

The primary goal of the Beta design studieswas to define a fully reusable HTOL mannedlaunch vehicle concept using near termtechnology. "Near term" meaning technologywhich would be mature enough forincorporation into a production vehicleassuming a ten-year vehicle development cyclestarting from the time of the studies. Near termlow risk technology dictated a two stage design(in order to meet vehicle propellant massfraction requirements) with a rocket secondstage (to avoid high risk undemonstratedscramjet propulsion). The design mission isshown in Figure 1. The vehicle was required tohave a self ferrying capability. That is, thebooster is capable of ferrying itself along with amated unfueled orbiter. The booster was to be

air-breathing with at least a sufficient amount oflow speed air-breathing engine thrust to meetthe ferry requirement. A bottom mounted

second stage was also desired for a morenatural separation procedure during staging, theheavier orbiter being dropped from the lightermore aerodynamic booster. As shown inFigure 2, this mounting technique also permitseasy ground mating operations with the orbitersimply being towed into place under thebooster.

Beta I

Beta I was conceived by the Air Force, WrightLaboratory, Flight Dynamics Directorate in1986 to meet the above design goals and acontract was awarded to the Boeing Space andDefense Group to analyze and refine the initial

design.1 It was sized to have a payloadcapability of 50,000 lbs. to low polar orbit. Thevehicle is depicted in Figures 3 and 4, which are3-view drawings of the orbiter and booster.

The orbiter is a lifting body design based on theX-24B experimental vehicle. The lifting bodydesign was adopted to meet an Air Forcerequirement for an orbiter with high cross rangecapability. The orbiter has a cold structuredesign consisting of a graphite/polyimidestructure and aluminum/lithium propellanttanks, with an external thermal protectionsystem similar to that of the Shuttle, but ofmore advanced and durable design. Thepropulsion system consisted of a singleLOX/LH fueled SSME rocket engine. Orbitergross weight is 600,000 lbs.

The booster is of a hot structure design withnon-integral propellant tankage. The structureconsists of titanium with Ren6 41 in the hot

areas and carbon-carbon leading edges. TheGLOW of Beta I is 1,900,000 lbs.

The booster propulsion system consists of twopodded over/under turboramjets, see Figure 5,and an SSME rocket. There were three

requirements which drove the design of thebooster propulsion system. A side-mountedpodded type design was required toaccommodate the bottom mounted orbiter.

Second, the self ferry requirement made itdesirable to use JP fuel for the low speed

2

propulsion system. Third, the desire to useexisting advanced design turbine engines forlow speed propulsion; the most advancedavailable at the time were those designed for theATF. The high speed propulsion system wouldconsist of LH fueled ramjets. The ATF enginesare too small to provide enough thrust for the

fully loaded Beta I vehicle during the launchmission with a reasonable number of engines,

so only enough engines to accommodate theferry mission were provided. The SSME, alongwith the orbiter SSME, provides the additionalthrust needed for low speed acceleration beforeramjet operational speed is reached. Propellantis cross fed from the booster to the orbiter

during the boost phase. The operating scheduleof the propulsion system during boosteroperation is shown in Figure 6. Staging occursat Mach 8. The orbiter SSME is not shut

down after the ramjets are fully operational,even though that would be more efficient,because the engine does not have a restart

capability.

Wind tunnel models of the booster and orbiterwere fabricated and tested from subsonic Mach

numbers to Mach 8. The models were tested

both in a mated condition and separately.Although the models were designed to perform

separation tests, these were never accomplisheddue to lack of funds. Despite the inability to do

the separation tests, the wind tunnel tests didprove the soundness of the aerodynamic designand the feasibility of flying the booster with theopen cavity on the underside after orbiterseparation.

A New Mission

Subsequent to the design of the Beta I, interestin the United States shifted to deploying

payloads of about 10,000 lbs. to low polar orbit(about 18,000 Ibs to the Space Station). In1990, NASA's Lewis Research Center becameinterested in examining the possibility ofdeveloping a near-term two-stage air-breathinglaunch vehicle to meet this requirement. The

Lewis design goals were very similar to thoseof the original Beta, with the exception of thepayload size. Therefore, it was determined that

Beta I would be an ideal starting point for thestudies. The main question that needed to beanswered was: "Could the vehicle be viablydownsized to one fifth of its payload size?"

Beta II

The Air Force was just finishing their contractwith Boeing when Lewis approached themabout downsizing Beta to deliver a 10,000 lbs.payload to low polar orbit. The Air Forceagreed to extend the contract for the downsizingstudy with Lewis funding. This gave Lewisaccess to the expertise which had beendeveloped at Boeing and at Wright Laboratory,thus greatly expediting and enhancing the Beta

II design study.

Upon examining Beta I in more detail, itappeared that the downsizing may afford the

possibility of enhancing Beta's operability andmaintainability even further. If the boosterrocket engine could be dispensed with, thenumber of propulsion system types on thebooster could be reduced, thus reducing spare

parts inventories and maintenance requirements.Historically, high performance liquid rocketshave been maintenance intensive.

Subsequent to the Beta I study, NASA andindustry initiated the HSCT research program.The projected service date for such an aircraftmatched very closely that required for thedevelopment of Beta. The turbine enginesbeing studied for the HSCT were roughlydouble the size of the ATF engines. This,combined with the downsizing of Beta, opened

up the possibility of designing a purely air-breathing booster propulsion system usingturbine engines which could be available fromanother source.

The evolution of the Beta launch vehicle,

beginning with Beta I, is shown in Figure 7.The downsizing process is shown by the firstfour boxes in the figure. Since the use of arocket during the boost phase was undesirable,the ramjets would be the only source ofpropulsion during the high speed portion of the

flight and, therefore, would have to operate at

3

full power in a high dynamic pressureenvironment. It was felt that Mach 6.5 would

be an upper limit for a practical conventionalramjet under these more severe temperature andpressure conditions. The first step in thedownsizing process was to reduce the stagingMach number to 6.5. The propulsion systemremained the same in this step. As would beexpected, the GLOW increased. Next, theturbine engine size was increased, the boosterrocket engine was removed and the orbiter

rocket was not fired. This new propulsionsystem schedule is shown in Figure 8. Thisexclusively air-breathing booster propulsionsystem design resulted in a 29% reduction ofGLOW, as shown in Figure 7. Finally, thepayload was reduced to 10,000 lbs. and thevehicle aerodynamics recalculated for thesmaller vehicle, resulting in an initial Beta IIGLOW of 1,200,000 Ibs. This vehicle is the

Interim Beta II listed in Figure 7. The weightincluded a conservative 20% booster growthmargin and a 10% second stage growth marginbased on stage empty weight. A larger growth

margin was included in the booster weight toaccount for the additional uncertainties in

estimating the weight of the air-breathing

propulsion system.

The interim Beta II booster is depicted inFigure 9. The general shape of Beta I wasretained at this point. The orbiter, which had agross weight of 345,000 lbs., is depicted inFigure 10. The orbiter was changed to a wing-body configuration because of the higherstructural efficiency of this type of airframe asopposed to the original lifting body design.Structural efficiency is more important in thesmaller Beta II orbiter and NASA did not have

the high cross range requirements of the AirForce. The SSME rocket engine propulsionsystem was retained for the orbiter.

The interim Beta II propulsion system is shownin Figure 11. It was an over/under designconsisting of 5 HSCT turbine engines mountedbelow a ramjet duct. Since turbine engines of afixed size were being used, enough engines hadto be installed to overcome transonic drag. Theinlet capture area was sized as a compromisebetween engine airflow requirements at high

speed and inlet transonic spillage drag

alleviation. The ramjet was mounted on top toallow access to the airframe for a large bypassduct to reduce the transonic spillage drag. Itwas recognized that the 5-turbine-enginearrangement was not ideal, but the initialcontract task did not have enough resources to

allow for further design iteration; therefore, thisdesign was adopted as an interim solution.Even though it was not possible to obtain amore optimum air-breathing propulsion systemdesign at the time, the design task did prove thefeasibility of the vehicle.

Subsequently, a more rigorous aerodynamicanalysis was undertaken. This analysisincluded reshaping the booster in an effort toimprove its L/D. The curves marked InterimBeta II and Fixed Inlet Beta II in Figure 12show the improvement in L/D that resultedfrom this effort.

A variable capture area inlet was incorporatedinto the propulsion system to further reduce

transonic spillage drag. The capture area is at aminimum during transonic flight and opens tofull capture at about Mach 4 and above. Figure13 depicts the minimum and maximum capturearea positions of the inlet. The upper surfaceof the nacelle is included in the vehicle

aerodynamics. Therefore, when the capturearea vanes the vehicle drag also varies, as can beseen in Figure 12 by comparing the curveslabeled Fixed Inlet Beta II and Final Beta Ii.

Although the vehicle drag increases in thetransonic region, it is more than compensatedfor by the reduction in inlet spillage drag andthe net effect is a positive one. The benefit of

the variable capture inlet can be seen in Figure14. The shorter acceleration time for the

variable inlet vehicle results in less fuel usageand therefore a lighter vehicle. The variablecapture inlet also removed the requirement for

the large bypass duct so that a moreconventional over/under layout of turbineengines on top and ramjet on bottom could beused as depicted in Figure 13.

Preliminary thermal management analysisresults indicated that it may not be possible tocool the ramjet without exceeding

4

stoichiometricfuel-air ratiosin theMach6 to6.5regime. It also appeared that the inlet mayrequire active cooling at these Mach numbers.Therefore, lower staging Mach numbers wereexamined. It was determined that reducing the

staging Mach number to 5.5 would result inonly a 5% growth in GLOW while significantlyreducing the risk in meeting the ramjet coolingrequirements. Even with this reduction instaging Mach number, the booster aerodynamicand propulsion refinement effort resulted in a17% reduction in GLOW from the interim

design (1,000,000 lbs. vs 1,200,000 lbs.), seeBeta II in Figure 7. The current Beta II boosterdesign is shown in Figure 15. The reduction in

staging Mach number resulted in an orbiterweight growth to 400,000 lbs. and an increasein _' from 0.811 to 0.827. Figure 16 displays

the weight breakdown of the booster and theorbiter.

Beta III

The JP/LH fuel combination on the booster

provides a very good compromise betweenvolumetric efficiency and high specific impulse.However, it does add the complexity of a dual

fuel system. Also, comments received fromlaunch vehicle operations people indicate thatLH fuel systems require high amounts ofmaintenance and operational activities.

Beta III, the final vehicle presented in Figure 7,has evolved from Beta II as a booster that is

fueled by JP only. Except for the fuel change,the propulsion system is the same over/underturborarnjet layout. JF' fueling of the ramjets ismade possible by the emerging endothermichydrocarbon fuel technology which increasesthe heat sink available in hydrocarbon fuels.Since this technology is in the early stages ofdevelopment, the Beta III propulsion systemhas a higher development risk than Beta II, butthe increased operational efficiency maywarrant the extra risk.

Beta III is in the early stages of design. The

staging Mach number is expected to remain at5.5. The GLOW is about 1,500,000 lbs. This

increase in GLOW is due to the lower energy

content of JP vs LH, higher system weight ofthe endothermic fuel cooling system and theaddition of two more turbine engines pernacelle to maintain a sufficient thrust to weight

ratio.

Operational Characteristics

Beta's principal mission is to provide flexiblelow cost access to space. The vehicle design is

a major step in that direction. The booster isessentially an airplane. No launch pads arerequired and stage mating can be accomplishedwith standard airplane towing equipment.Vehicle processing would be done in airplane

hangar-like facilities. In addition, it isenvisioned that payload processing will occuroff line from orbiter processing with standard

payload interfaces being provided by the orbiterto which all payloads would conform.

An advantage of the two-stage air-breathingbooster design is the possibility of performingoffset launches. In this type of launch profile,the booster cruises for up to several hundredmiles before going into the launch trajectory; ineffect, changing the point of launch relative tothe earth's surface. This makes possible onebase operations for both polar and easterlylaunches from the United States, greatly

enlarged launch windows and satelliterendezvous within one and one-half orbits.2

One base operation from Kennedy SpaceCenter can be accomplished by using the offsetlaunch capability to perform polar launches, seeFigure 17. The vehicle would takeoff fromKennedy, fly a subsonic cruise across Floridaa.nd then turn south in the Gulf Of Mexico to

go into the polar launch trajectory. The orbiteris essentially at orbital flight conditions beforeencountering any land mass and the booster cancomplete its turn for the return flight toKennedy within the confines of the Gulf.

Another key aspect to Beta's cost effectivenessis its versatility. Once the booster is developed,it would be capable of carrying other types ofpayloads besides the orbiter. Somepossibilities are shown in Figure 18. A primaryexpansion of capability could be accomplished

5

by developing an expendable upper stage.Developing such an upper stage would allow

the placement of medium weight payloads intoorbit, Figure 19. The expendable upper stagemay also be applicable, perhaps withmodifications, as an upper stage for a futureexpendable heavy lift launch vehicle.

An experimental hypersonic research aircraftcould also be carried by the Beta booster. Thiswould permit an evolutionary research path toattaining the ultimate goal of an SSTO orhypersonic cruise vehicle. As depicted on theleft side of Figure 20, a hypersonic researchvehicle which is dropped from Beta would notneed a low speed propulsion system andtherefore would be less risky and costly todevelop. The research aircraft developmentcould concentrate on its main mission, that is, to

investigate hypersonic flight. Also, as shownon the right in the figure, Beta would stillremain an important launch vehicle even afterthe development of an SSTO. In addition toproviding a back-up to the SSTO, with theexpendable upper stage it would remain themain vehicle for launching medium sizepayloads.

Finally, once the booster is developed, it couldalso be used as a carrier to launch a high speedcargo transport aircraft, which would cruise atabout Mach 5. Because the cargo transportpropulsion system could be of much simplerand lighter design than one required to operatefrom takeoff to the cruise Mach number, a

longer range and/or larger payload could beachieved than for a single aircraft. In thisscenario, there would be a Beta boosterstationed at each end of the route.

Conclusion

Beta is a viable and robust air'breathing twostage horizontal takeoff and landing launchvehicle. It is of conservative design with largegrowth margins included. Beta II/III are in aweight class similar to advanced versions of the747 and the Russian AN225.

Beta is designed for low cost airplane-like

operations. It can take off and land from

standard Strategic Air Command runways andis fully recoverable and reusable. Stage matingcan be accomplished with standard airportequipment and the vehicle is self ferryable.

Beta can provide a near term solution to lowercost manned access to space and Space StationFreedom support. With the inclusion ofexpendable upper stages, Beta II/III has thecapability to place up to 45,000 lbs. into loweasterly orbit.

Finally, development of a TSTO launch vehiclesuch as Beta would permit an orderly and

evolutionary progression to the development ofair-breathing SSTO and hypersonic cruisevehicles.

.

o

References

Wetzel, F_, D.; Meadowcroft, E. T.; Paris, S.W.; Dunn, B. M.; Weldon, V. A.; Kotker,D. J.; Williams, D. P.:Research Vehicle

Configurations For Hypervelocity VehicleTechnology. WRDC-TR-90-3003,Volumes I and II, Apr. I990.

Lt. Col. Alan J. Parrington, USAF: TowardA Rational Space-TransportationArchitecture: Airpower Journal, Winter1991, Pages 47-62

6

Figure 1. Beta Design Mission

Figure 2. Mating Operations

7

152' "1

GLOW = 600 K lb.

Payload = 50 K Low Polar

Growth Mar. = 2%

1' = 0.805

Matenals

Graphite PolyimideCeramic Tiles (TPS)Aluminum/Lithium (Cryo Tanks)

Propulsion1 SSMELOX/LH RCS

60' ---_-_1

Figure 3. Beta I Orbiter Concept

I

I__ 280'

/]

rl

GLOW = 1.9 M lb.

EW = 4.39 K Ib.

Growth Margin = 2% EW

1'= 0.66

MaterialsTitaniumRen6 44Carbon-Carbon (leading edges)

Propulsion8 ATF turbofans (JP)2 Ramjets (LH)I SSME Rocket (LOX, LH)

Stages M = 8

195'

Figure 4. Beta I Booster Concept

8

Turbofan

M=Oto3

and Ramjet Operation

M=3to8

Ramjet Operation

Figure 5. Beta I Nacelle Concept

Full Power

Partial Power

Orbiter Rocket

Booster Rocket

Booster Turbojet

Booster Ramjet

Figure 6.

0 2 L

M

Beta I Engine Operating Schedule

9

6 8

2.0-

1.5-

TOGW

106 Ibs.1.0-

/

0

• I00% Airbrazt_g

Figure 7. Evolution of Beta Airbreathing Launch Vehicle

Full Power

Partial Power

Orbiter Rocket

Booster Turbojet

Booster Ramjet

0 2 4 6M

Figure 8. Beta II Engine Operating Schedule

10

7

/71

GLOW = 1.2 M Ib.

EW = 490 K lbs.

Z ' = 0.436

Materials

Reng 41

Inconel 718

Carbon-Carbon (leading edges)

Propulsion

10 HSC T Turbine Eng.

2 Ramjets

Stages M = 6.5

m

Figure 9. Interim Beta II Booster Concept

< 67' 2"

T23' 4"

IL

GLOW = 345 KIb.

Payload = 10 K Ib. (Iow polar)

_.' = 0.811

Materials

Graphite Polyimide

Ceramic Tiles (TPS)A/uminundLithium (Cryo tanks)

Propulsion1 SSME

2 RLIO

LOX/I_H RCS

Figure 10. Mach 6.5 Staging Beta II Orbiter Concept

11

M=Oto3

Turbofan and Ramjet Operation

M=3to6.5

Ramjet Operation

Figure 11. Interim Beta II Nacelle

6

5

BOOSTER 4

LIFT-TO-

DRAG3

RATIO

2

0

\\

IIIi

0 1

FIXED INLET BETA IIFINAL (VARIABLE INLET) BETA II

o

INTERIM BETA II

I , I I _, T 1

2 3 4 5 6 7

MACH NUMBER

8

Figure 12. Beta ii Lift to Drag Comparison

12

Math No. = 0 to 3 4 HSCTTurbojets lAB

Thrust ,, 80K Ibs. / eng.Fuel - JP7

v ,-, _ _ RamjetFuel - (Beta II) LH2

- (Beta III) JP7

Mach No. = 3to5.5

Ramjet Operation

i

Figure 13. Beta II Nacelle Concept

7

6

5

MACH 4 /

NUMBER VARIABLE INLET BETA II /

3 TSEP= 493 SEC /

2 /i

JJ

//

/

/

FIXED INLET BETA IITSEP = 592 SEC

O T

0 100 200 300 400 500

TIME (SEC)

600 700 800

Figure 14. Inlet Study Trajectory

13

GLOW = 1.0 M Ib.

EW = 440 K Ib.

Growth Mar. = 20% E.W.

Z ' = 0.31

Mateda/s

Ren_ 41

/cone/718

Propulsion8 HSCT turbojets

4 LH ramjets

Stages M = 5.5

255" 142"

Figure 15. Beta II Booster Concept

Booster Orbiter

t ,200,000

900,000

600,000

300,000

(O¢olter)4O2,337Ibs.

Propelanls199,112 Ibs.

Margin 73,509 Ibs.

Entrywe_t366,599 R:_,

Crew, Reslcluols& Reserves3.349 Ibs.

/

440

4OO

360

320

280

K Ibs. 240

20O

160

120

8O

40

0

propelant(,Ascent& Orbit)324,525Ibs,

Err,p_,We_gt_t58,915 Ibs

_----- Paytoad 10,000Ibs.

Crew + Residuals3,006G_owlh 5,891 Io_.

Figure 16. Beta II Weight Breakdown

14

Polar Launch From KSC

Subsonicover

Florida

Figure 17. Single Site Operation

Reusable Manned Scramjet Research Vehicle(x- 30)

Expendable Upper Stages

JMilitary Cargo Transport

(M = 5 Cruise)

Figure 18. Beta II Versatility15

20 K ib Easterly10 K lb + 10 Crew SSF

f \ 10 K Ib + 2 Crew Polar

\\

\ \

Beta

20 K Ib GTO

45 K Ib SSF

40 K Ib Polar

\

Figure 19. Year 2005 Launch Vehicle Architecture

2005 2020

20 K ib Easterly 45 K Ibs SSF10 K Ib Polar 20 K Ibs GTO

f _ SSF Supply / Crew Rotation 40 K Ibs Polar

Beta Carder for X30 SSTO Beta

Figure 20. Beta Enhances SSTO Development

16

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December 1992 Technical Memorandum

4. TITLE AND SUBTITLE

The Design and Evolution of the Beta Two-Stage-to-Orbit HorizontalTakeoff and Landing Launch System

6. AUTHOR(S)

Leo A. Burkardt and Rick B. Norris

7. PERFORMING ORGAN_m_TION NAME(S) AND ADDRESS(ES) " --

National Aeronautics and Space Administratio¢Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space AdministrationWashington, D.C. 20546-0001

1'1. SUPPLEMENTARY NOTES

5. FUNDING NUMBERS

WU-505--69--40

PERFORMING ORGANIZA_ONREPORT NUMBER

E-7776

10. SPONSORING/MONITORINGAGENCY REPORTNUMBER

NASA TM-106118AIAA-92-5080

Prepared for the Fourth International Aerospace Planes Conference sponsored by the American Institute of Aeronautics and

Astronautics, Orlando, Florida, December I---4, 1992. Leo A. Burkardt, NASA Lewis Research Center, Cleveland, Ohio. Rick B.

Norris, Wright Laboratory, Wright-Patterson Air Force Base, Ohio 45433. Responsible person, Leo A. Burkardt, (216) 433-7021.

f 2ao DISTRIBUTION�AVAILABILITY STATEMENT

Unclassified -Unlimited

Subject Category 15

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13. ABSTRACT (Maximum 200 words)

The Beta launch system was originally conceived in 1986 as a horizontal takeoff and landing, fully

reusable, two-stage-to-orbit, manned launch vehicle to replace the Shuttle. It was to be capable ofdelivering a 50,000 lb. payload to low polar orbit. The booster propulsion system consisted of JP fueledturbojets and LH fueled ramjets mounted in pods in an over/under arrangement, and a single LOX/LHfueled SSME rocket. The second stage orbiter, which staged at Mach 8, was powered by an SSME rocket.

A major goal was to develop a vehicle design consistent with near term technology. The vehicle designwas completed with a GLOW of approximately 2,000,000 Ibs. All design goals were met. Since then,interest has shifted to the 10,000 lbs. to low polar orbit payload class. The original Beta was down-sized tomeet this payload class. The GLOW of the down-sized vehicle was approximately 1,000,000 Ibs. Thebooster was converted to exclusively air-breathing operation. Because the booster depends on conventionalair-breathing propulsion only, the staging Mach number was reduced to 5.5. The orbiter remains an SSME

rocket-powered stage.

14. SUBJECT TERMS

Hypersonic; Two-stage-to-orbit; Launch vehicles; Reusable

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