8
Acto Astronautica Vol. 37, pp. 87-94. 1995 Elsevier Science Ltd. Printed in Great Britain SINGLR-!3TAGE-TO-ORBIT - A STEP CLOSRR D&MC. Freeman, Jr.*, Douglas 0. Stanley** CharlesJ.Camarda***andRog~ALegsch** NASA Langley Reseamlt Center Hampton, Virginia 23W-OBBl Stephen A. Cook**** Marshal1space Flight Center Huntsville. Alabama 35812 ABSTRACT Over the past several years there has been a significant effort within the United States to assess options to replace the Space Shuttle some time after the turn of the century. In order to provide a range of technology options, a wide variety of vehicle types and propulsion systems have been examined. These vehicle concepts which are representative of the classes of concepts that could be pmposed for any future vehicle development is being used in the initial phase of the access to space activity to identify rquirements for the technology matumtion effort and to assess approaches to achieve the required low operations cost. This paper provides the results of recent systems analyses and describes the ongoing technol- ogy matumtlon and demonstration program supporting the Reusable Launch Vehicle Program. NOMENCLATURE ACC advanced carbon-carbon Al-U aluminum lithium ALT advanced launch technology APU auxiliary power unit %ader, Spe Tnanspottatt’on Q@ice +*Aemepnce Bngineer, Space Systems & Concepts Division **‘Wd Thermal Smtcnues Bnmch ****Systems Engineer, Launch Vehicle Technology @lice This p8pa is declared a work of the U.S. Oovemment and is not subject to copyright protection in the United States. ARC ATS BITE CFBI CJSiC DOD ECD GPS IIUI ISSA K!IC L&C LH2 LO2 LRU Me MMC MSFC NDE Ree RLV Rn SIP SSTO STS TAB1 Ti-Al TPS TRL X Ames Research Center access to space built-in test equipment composite flexible blanket instdatlon carbo&ilicon carbide DepWment of Defense ele&caI conversion and distribution global positioning satellite internal multi-screen insulation International space station Alpha RemudyspaceCenter Langley ReseamhCenter liquid hydrogen (at 4.42 lWf$) liquid oxygen (at 71.14 Wft ) line-replaceable unit Mach number at boum%y layer edge metal-matrix composite Mar&all space Flight Center nondestructive evahtation momentum-thickness Reynolds number reusablelaunchvehicle nose radius stralnisolationpad single stage to orbit space won system tallorable advanced blanket htsuRtion titaniumaluminide thermal promotion system technology readiness level vehicle lmalth vt VedCRlml verticdtakmfUv&~ clistamfromnose0fvellicl~h 87

Single-stage-to-orbit — A step closer

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Acto Astronautica Vol. 37, pp. 87-94. 1995 Elsevier Science Ltd. Printed in Great Britain

SINGLR-!3TAGE-TO-ORBIT - A STEP CLOSRR

D&MC. Freeman, Jr.*, Douglas 0. Stanley** CharlesJ.Camarda***andRog~ALegsch**

NASA Langley Reseamlt Center Hampton, Virginia 23W-OBBl

Stephen A. Cook****

Marshal1 space Flight Center Huntsville. Alabama 35812

ABSTRACT

Over the past several years there has been a significant effort within the United States to assess options to replace the Space Shuttle some time after the turn of the century. In order to provide a range of technology options, a wide variety of vehicle types and propulsion systems have been examined. These vehicle concepts which are representative of the classes of concepts that could be pmposed for any future vehicle development is being used in the initial phase of the access to space activity to identify rquirements for the technology matumtion effort and to assess approaches to achieve the required low operations cost. This paper provides the results of recent systems analyses and describes the ongoing technol- ogy matumtlon and demonstration program supporting the Reusable Launch Vehicle Program.

NOMENCLATURE

ACC advanced carbon-carbon Al-U aluminum lithium ALT advanced launch technology APU auxiliary power unit

%ader, Spe Tnanspottatt’on Q@ice +*Aemepnce Bngineer, Space Systems &

Concepts Division

**‘Wd Thermal Smtcnues Bnmch ****Systems Engineer, Launch Vehicle

Technology @lice This p8pa is declared a work of the U.S. Oovemment and is not subject to copyright protection in the United States.

ARC ATS BITE CFBI CJSiC DOD ECD GPS IIUI ISSA K!IC L&C

LH2 LO2 LRU

Me MMC MSFC NDE

Ree RLV

Rn SIP SSTO STS TAB1 Ti-Al TPS TRL

X

Ames Research Center access to space built-in test equipment composite flexible blanket instdatlon carbo&ilicon carbide DepWment of Defense ele&caI conversion and distribution global positioning satellite internal multi-screen insulation International space station Alpha RemudyspaceCenter Langley ReseamhCenter liquid hydrogen (at 4.42 lWf$) liquid oxygen (at 71.14 Wft ) line-replaceable unit Mach number at boum%y layer edge metal-matrix composite Mar&all space Flight Center nondestructive evahtation momentum-thickness Reynolds number

reusablelaunchvehicle nose radius stralnisolationpad single stage to orbit space won system tallorable advanced blanket htsuRtion titaniumaluminide thermal promotion system technology readiness level vehicle lmalth vt VedCRlml

verticdtakmfUv&~

clistamfromnose0fvellicl~h

87

88 45th IAF Congress

INTRODUCTION

Over the past several years there has been a significant effort within the United States to assess options to replace the Space Shuttle some time after the turn of the century. l-3 In order to provide a range of schedule and technology options, a wide variety of vehicle types and propulsion systems have been examined. These include single-stage and two-stage systems, systems utilizing rocket and airbreathing propulsion,

systems for personnel and/or cargo transportation, and sys- tems with varying degrees of reusability.4‘6 In January 1994 the NASA Access to Space (ATS) study was completed (see Figure 1) with a recommendation for the development of a new fully reusable system to replace the Space Shuttle in the 2005 to 2010 time period (Option 3). Within the past several

months there has been an increased interest in creating a National program to address access-to-space issues which has culminated in a National Space Transportation Policy issued by the White House Office of Science and Technology Policy in August 1994. This policy designates NASA as “the lead agency for technology development and demonstration of next-generation reusable space transportation systems.” With this direction NASA has focused the current Reusable Launch Vehicle (RLV) Program on assessing a number of vehicle concepts, maturing the required technologies to an acceptable level, and identifying candidate approaches to demonstrate the performance and operability of these technologies for a reus- able flight vehicle.

Providing a space transportation system with low-cost operations is the major overriding requirement for the current Reusable LaunchVehicle Program. A rocket-powered single- stage-to-orbit (SSTO) vehicle has been chosen as the initial goal for the technology and flight-demonstration activities.‘l

Aceosa to Spr~ea Vision

Figure 1. Access to Space Study Description.

The current activity is focused on a number of concepts proposed by both government and industry (see Figure 2). Included in these concepts are winged and lifting-body verti-

cal-takeoff/horizontal-landing (VTHL) configurations and a conical vertical-takeoff/vertical-landing (VTVL) concept. These vehicle concepts (which are viewed as being represen- tative of the classes of concepts that could be proposed for any future vehicle development) are being used in the initial phase of the program to identify requirements for the technology

effort and to assess approaches to achieve the required low operations cost. This paper will present the results of recent system analyses to identify those technologies required to meet the SST0 performance, operability, and reusability goals with acceptable risk. ‘Ihe ongoing and planned technol- ogy demonstration program will be described with relevance

shown to the development of any future operational system

Conical Powered-Larder VNL

Figure 2. Representative Reusable Launch Vehicle Study Concepts.

SST0 REFERENCE MISSION AND CONFIGURATION

As a part of the Access to Space Study, a reference mission was developed to capture a very large portion of future U.S. civil. military, and commercial payloads. Although there is an ongoing effort to define mission requirements for the RLV Program, this mission model is currently being used to design reference vehicle concepts like the VTHL SST0 ve- hicle shown in Figure 3. The reference SST0 vehicle in Figure 3 would deliver to the International Space Station Alpha (ISSA)andreturna25-klbpayloadwitboutcrewwhenIaunched from the Eastern Test Range at the Kennedy Space Center (KSC). The ISSA will be located in a 220-nmi circular orbit inclined 51.6 degrees to the equator. Four personnel, consumables, and refrigerated storage lockers could be ac- commodated in a pressurized ISSA crew rotation module located in the forward portion of the payload bay. This same

4Srh IAF Congress 89

module would also be used, with minor modifications, for satellite servicing missions. The vehicle is designed to be flown in an unmamted mode. ‘Ihe payload bay is lS-ft. in diameter and 30- A. long. Onboatd propehant would provide an incremental velocity of 1100 ft/sec following launch inser- tion into a SO by 1MMmi orbit. Landing would nominally be at the KSC launch site. The SST0 has a 11tBnmi crossrange capability to allow once-around abort for launch to a polar orbit and to increase daily landing oppottunities to selected landing sites. lbe SST0 also has a large range of intact abort oppottunities in the event of a forced shutdown of a single main engine.* Passenger escape is provided by ejection seats in the appropriate portions of the flight regime.

Figure 3. Winged VTHL SST0 Configuration Used for Technology Trade Studies.

Although there are several vehicle options, many of the tmdestudies have been made with areference vehicle concept. This reference vehicle (presented in Figure 3) is a vertical- takeoff, horizontal-landing winged concept with a circular- cross-section fuselage for structural efficiency. The payload bay is located between an aft liquid hydrogen (LH2) tank and a forward liquid oxygen &02) tank. The normal-boiling- point LH2 and LO2 propellants are contained in integral, reusable cryogenic tanks. Two cylindrical kerosene fuel tanks are located underneath the payload bay. The SST0 main prop&ion system uses seven tripropellant engines to lower system dry weight. The vehicle employs wing tip fms for directional control rather than a single vertical tai1.9 The vehicle employs a smndardixed payload canister concept with common interfaces to allow off-line processing of payloads andrapidpayload integration. Evolutionary propulsion, struc- ture, thermal protection system, and subsystem technologies are utilized that are consistent with an iei9al opemting capa- bility by 2008.

TECHNOLOGY PAYOPP

AspartoftheinitialAccesstoSpacesadies,aneffortwas made using the reference concept to quantify the potential performance benefit from incorporating advanced technolo- gies into candidate SST0 concepts. The analyses were con- ducted using Space Shuttle technologies as the point of departure and then assessing the additive benefits of each of the advanced technologies for SST0 applications. These results are presented in Figures 4 and 5. The differences between proposed SST0 technologies and Space Shuttle technologies to provide a point of departmeam SUllUIltized

in Table 1 .l” The effect on vehicle dry weight of the technol- ogy improvements that were identified in the Access to Space Study is presented in Figure 4. The analyses show that, if a SST0 vehicle weredesigned with Space Shuttle technologies for the RLV mission, the vehicle would weigh 552.000 lb. If advanced technology subsystems were utilized and the ve- hicle were resixed as shown in Figure 5, the weight reduction

would be 3 1,000 lb, and the vehicle would weigh 5 15,ooO lb. Likewise,usinganadvancedthermalprotectionsystem(TPS) and adding aluminum lithium tanks would further reduce the weight for the resixed vehicle to 379,000 lb. Adding a com- posite primary structure (non-pressurixed. load-bearing struc- ture) and a composite hydrogen main propellant tank has a significant impact on the vehicle dry weight, resulting in a resixed vehicle that weighs 206,tMO lb. With the accumulated weight benefits of the advanced subsystems, advanced TPS, composite structure, and composite tank, a SST0 vehicle with an estimated dry weight of 206,000 lb can be achieved. The weight and size of this vehicle would be small enough to reduce development risk to an acceptable level and to provide a reasonable-size vehicle to meet the operability goals of the program. As shown in the figure, additional weight reduc- tions can be achieved through the use of advanced propulsion and materials technology. The ATS study concluded that, with a reasonable technology maturation effort, SST0 may be achievable in the timeframe projected for a new, reusable launch vehicle by the recently issued National Space Launch Strategy.

90 45th IAF Congress

Table 1. SST0 Enabling and Enhancing Technologies.

cammlscnluioll

OhWRCS pop&ion

Rimpowcr

ESD system

AVionica

Gtwnd opcmtions

nit opuwions

t3qbitc composite. honeycomb WithhgfrmwS

Rcuubk, ALLi of papbite composite skin-attingu with frame4

Graphite composite and Al-U. conical ormJss

ACC or CfSiC hot structure

ACC or USE hot stmcture

Reusable. extemeJ or internal , bonded, closed-all foam or lh4I

Bonded flexible blankets. ceramic tiles, or mechanically attached panels

Single or dual t&l, hydrostatic bearings. pneumatics aud EMA. VHM. oxygen- rich pt&tuwr, g&gas main injection. simplificdttnbotnachinuy, jet boost pumps, altkude-wmpcnsating nozzle. advanced rm&r*la. composite feed l&s

Rmmable, T&U. stainless steel. and graphhecomposite.EMA

EMA

High-power density fuel cells. 27OMcIBolrW each

Pibemp4in and copper wirr. 270VDC

Fiberoptics* flat panel displays. GPS rccciww. beads-up display, ring laser gyms.

Mmicmpmccrson. SUtUtSUlUXS.Vlfhl

srre, LRU, Stawhbd payload cudNus.k&-ftecjobtteMdsuls. multiple pl kak detection. automated umbiiiuqmvcddiscamas. non-pym tclee? me&a&m8

VHM,ada@tiveguid8occ,~0IMtcd softwatepuJtomatcd misniat plmmbtg, automwcd tadewus ~docldnl.~detanuMgunmt. fault-tolmtlt uchik8xlues

Al~sldtl-m withrittgfrsttm

Jzxpdsble, ablminum skin-StIlngu with tiatncs

Titanium bom&poxy. and ahtmittttmtNss

AlumbIumwithbonde&TPs

Gubon~nhotstmctuu

Expendable. external. spray-on closed-cell foam

Bon&d ceramic tiles and flexible

Space Shuttle Main Engine. hydmtdiis and pneumatics. stainless steel and ahuninum feed lines

&e-fall. stainless steel and aluminum hydtwlics

Hydraulics

Integrated hypergolic

Hydmziac APU. fuel cells. 28VDCfl2kW each

Copper wire. 28 VDC

Copper wire. cathode-my tube

divw% nwncdc op tmss -tY

STS opemtia

STS operations

45th IAF Congress 91

100% 21511

20% 106tt

78% 167R

70% 151ft

POybad=45,000 Ibto100rwnl.28eottM $& 500

500

400

300

200

100

0

- Cumulative lncorpomtion of advaned teehnotogiea -

Figure 4. Effect of Technology on Vehicle Dry Weight for the Winged VTl% SST0 Concept.

Figure 5. Effect of Technology on Vehicle Size for the Winged VTHL SST0 Concept.

IftbetechnologyberlefitspresentedhtFQure4prove to becorrect,thedryweightreductionfequitdtoxneetSsM petfwce can easily be achieved. Additional weight teduc- tions can be utilized as margin to reduce vehicle development risk as shown in Figure 6. Using a combination of composite. structure, composite tanks, and rripropellent propulsion pro-

vides a 31- percent dry-weight margin for a SST0 with a 233.OOMb dry weight. In addition to reducing tk develop- ment risk, a portion of this margin can be used to address the other major challenge of the RLV Program-to significantly reducetheoperatingcost.Forexample,someofthisadditional dry weight savings could be used to incorporate robusmess into the design of the vehicle subsystems to improve operabil- ity. The challenge for the activity currently being initiated in the RLV program is to quantify and develop metrics that not only address the SST0 performance requirements but also address the technology and engineeting design fern to assure that the vehicle is reliable, maintainable, and support- able in much the same manner as modern day a&raft. The RLVProgmmiscummtlyproceedingwithactivitiestomatute the required technologies leading to a flight-demonstration programtoassurethatanynewRLvdoesmeetthe10w- recurring-cost requirement. The technology matum&m and flight-demonstration programs wiI1 provide the data to veri& theperfotmanceassumptionsusedintheinitialcoxe+aI analyses aad provide an understanding of how opc&ility improvements af&ct vehicle weight and pexformance. llten, anin6nmeddecisioacanbema&conceAngtheviabilityof an operational SST0 in the RLV program timeframe.

45th IAF Congress

Vehkl@ dry wlghl groMh margin

. Paybd I 45,Wg lb to 100 nmi. 28’ orbit l Vehidr dry Might constant 0 233 KIb

- Cumuhtlvo Incorpadon of advmtwd teahnologks A

Figure 6. Effect of Technology on Vehicle Weight Growth Margin.

Figure 7. Technology Maturation Impact on Vehicle Devektpment Cost and Schedule.

The initial efforts in the fotmulation of a program to support the development of a future space transportation system are focusing on maturing rocket propulsion, composite and ahmritmm lithium structures, and metallic and ceramic TP!3 technologies to meet the SST0 performance and oper-

ability tcq uimtnents. The benefit of advanced technology matur&m &ivity for the RLV Program is pnsented in Figure 7. This fgure shows the relationship of SSTG vehicle development cost &td schedule to the technology readiis level (TRL.) at program inception. Tbe goal of the RLV technology program is to mature all required technologies to

TRL 6 (see Figure 8). These results show a 45% reduction in the vehicle development cost if the technologies are matured to TRL 6 in advance. The technology program will focus on maturing technology options that have been analyzed in the ATS study. The RLV F’rogratn is using cooperative NASA/ industry resources to mature the technology to an acceptable level of risk (I’RL 6) before proceeding with a vehicle devel- opment program. The technology program will also provide initial test results to verify the structural weight and propulsion characteristics utilized in the ATS study, which indicated the

viability of SST0 for RLV applications.

Figure 8. Technology Readiness Level Definition.

45th IAF Congress 93

TECHNOLOGY MATURATION PROGRAM

Since completion of the ATS Study, NASA has begun

developing a program for the technology maturation, flight demonstration, and development of a RLV. This program is being implemented with the NASA Marshall Space Flight Center @WC) being the lead center. The Advanced Launch Technology (ALT) Program was established in January 1994 to address the technology maturation and demonstration for the RLV Development Program. The ALT Program is focused

on maturing the technology for the RLV Program as shown in Fiiures 9 and 10. The technology maturation activity cur- rently in place and its relationship to the reference vehicle concepts is shown in Figure 9. The proposed schedule for the technology maturation of tbe cryogenic tank, composite pri-

mary structure, TPS, and advanced propulsion systems is shown in Figure 10. The large-scale demonstrations scheduled in the 1997-98 time period for these maturation activities are also shown in Figure 10. Some of the hardware developed and

Figure 9. Advanced Launch Technology Program.

Figure 10. Advanced Launch Technology Program Schedule.

fabricated in the technology maturadon program will be flight-

tested on the modified DC-XA vehicle from the Deprrrtmmt of Defense (DOD) Delta Clipper Program. Cutrent planning also

provides for a new advanced technology demonsttator as part of the development of any new RLV. The technology mamration is being achieved thmugh a NASA/industry coopemdve agree- ment. This cooperative NASA/ii activity was initiated in June of 1994 to address the enabling technologies identified by the ATS Study. The cooperative agreements am focused on

manning critical technologies that are common to all of the concepts being studied. The cooperative agmements focused on the major vehicle technology efforts am summa&ed below. Additional information concerning the complete set of technolo- gies required for SS’KI development is contained in Reference 10.

Reuspb)e. For years the development of a reusable cryogenic tank has been considered as the most difficult development challenge for a RLV. The development of tank and insulation materials and the design and manufac- turing of large-scale, flight-weight tanks are a significant challenge for the RLV Program. The thermo-structuraJ con- cept for either an aluminum lithium or graphite composite cryogenic tank that can support SST0 performance require- ments and still be readily inspected and repaired is a signifi- cant challenge. This is very evident from the schematic of a candidate thermo-structural approach presented in Figure 11.

The complexity of tank/insulation/TPS concepts is shown in Figure 11. These concepts can be as much as Cinches thick with several different materials and structural interfaces. A successful reusable cryogenic tank technology program is essential to enable a SST0 development. This program will include fabrication and testing of small-scale test articles and

large-scale pressurized test articles. Presented in Figure 12 is a schematic showing several thermo-structural concepts that are being studied as a part of the RLV Rogram. These concepts are presented to illustrate the complexity of the structure to accommodate cryogenic internal temperatures and the ascent and entry aerodynamic heating on the outer surfaces.

Presented in Figures 13(a) and 13(b) are descriptions of the specific tasks being performed in the cooperative agree- ments to address the complex cryogenic tank design problem and mature technology for development of both aluminum lithium and composite tanks. This task is being managed by

the NASA Marshall Space Flight Center (MSFC) with support from the NASA Ames Research Center (ARC) for the ceramic TPS and from the NASA Langley Research Center (LaRC) in graphite-composite primary structure development and ther- mal/structural analysis and testing. The tasks described in

94 45th IAF Congress

safety,anddecreaseddevelopmentando rationalrisk,thereby 6. Stanley,D. O.,Engelund. W.C.. Wi1hite.A. W.,andLaube, leading to a more affordable system. 18” J.,“‘AComparisonofSingle-StageandTwo-StageAirbm&ing

Launch Vehicles at High Staging Mach Numbers,” Journal of REFERENCES Soacecraff, Vol. 29, No. 5, 1992, pp. 735-740.

1. Freeman, D. C.. Wilhite, A. W., and Talay. T. A., 7. Stanley, D. 0.. Engelund. W.C.,Lepsch, R. A., McMillin, “Advanced Manned Launch System Study Status,” LAFPaper M.. Wurster, K. E.. Powell, R. W., Guinta, A. A., and Unal R., 91-193. Oct. 1991. “Rocket-PoweredSingle-Stage Vehicle Configuration Selec-

tion and Design,” AIAA Paper 93-1053, Feb. 1993. 2. Stanley, D. O., Talay, T. A., Lepsch, R. A., Morris, W. D.,

and Wurster, K. E., “Conceptual Design of a Fully Reusable 8. Nafte.1, J. C.. Engelund, W. E., Lepsch, R. A., Powell, R. W., Manned Launch System,” Journal, and Bacon, J.. “Ascent Abort Capability for a Rocket-Pow- Vol. 29. No. 4,1992, pp. 529-537. ered Single Stage Advanced Manned Launch System,” AJAA

Paper 93-3694, Aug. 1993. 3. Stone, H. W. and Piland, W. M., “An Advanced Manned Launch System Concept,” IAF Paper 92-0870, Aug. 1992. 9. Powell, R. W. and Freeman, D. C.. Jr., “Application of a

Tip-Fin Controller to the Shuttle Orbiter for Improved Yaw 4. Piland,W.M.andTalay,T.A..“AdvancedMannedLaunch Control,” Jo-, Vol. System Comparisons,” JAF Paper 89-221, Oct. 1989. 5. No. 4, 1982. pp. 325-329.

5. Freeman, D. C., TaJay, T. A., Stanley, D. O., and Wilhite, 10. Stanley, D. 0. and Piland, W. M., “Technology Rquire-

A. W.. “Design Options for Advanced Manned Launch Sys- ments for Affordable Single-Stage Rocket Launch Vehicles,”

terns.” AJAA Paper 90-3816, Sept. 1990. IAF Paper 93-V.4627, Oct. 1993.