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AIAA 93-4160 The Next Generation Manned Launch System - A Complex Decision Kathryn E. Wurster, Lawrence F. Rowel1 NASA Langley Research Center Hampton, VA Lewis L. Peach, Jr. NASA Headquarters, Code DD Washington, DC

AIAA Space Programs and Technologies Conference September 21 -23,1993

Huntsville, Alabama ‘d

For permission to copy or republish, contact the American lnstitue of Aeronautics and Astronautics 370 L’Enfant Promenade, S.W., Washington, D.C. 20024

THE NEXT GENERATION MMhED LAUNCH SYSTEM - A COMPLEX DECISION

Kathqn E Wurster’, Lawrence F Rowell** W NASA Langley Research Center

Hampton, VA

Lewis L Peach, Jr *** NASA Headquarters. Code DD

Wasiungton. DC

ABSTRACT

Continued mannedaccess to space will soon require upgrade orreplacement of the current SpaceTransportation System. Many advanced manned launch concepts have been proposed and an- alyzed to gain a better understanding ofthe technical issues, op- erational complexities. technology requirements, and life cycle costs. Performance goals must be traded against operational con- siderations and an architecture selected which captures the mk- sion requirements yet conforms to the budgetary constraints necessarily placed on the next generation of vehicles.

This paper addresses the critical issues that must be con- sidered in the selection of the next generation space rransporta- tion system. Mission requirements, design approach, technology level assumptions, and costconsiderations, are shown to strongly influence the type of system that will be favored. Over the past 15 years NASA, in cooperative efforts with industry. has con- ducted a number ofstudies regarding possible replacement strat- egies for the current system. This paper draws on the results of those systems studies to illustrate several vehicle trades which are discussed with respect to these issues. The trades include: reusable vs. expendable, rocket vs. airbreather, single vs. two- stage. hybrid airbreatherhocket systems. horizontal vs. vettical launch, and air-launch options. Three basic approaches to the next generation transportation system are identified: Shuttle evolution, an expendable-based architecture complemented by a reusable payloadlpersonnel Carrier, or a “clean sheet approach” based on advanced technology.

NOMENCLATURE

AB airbreather AI-Li aluminum lithium DC-X Delta Clipper Experimental

*Aerospace Engineer, Vehicle Analysis Branch, Space Sysrems Division. **Assisrant Head, Vehicle Analysis Branch. Member AIAA. **“Director, Advanced Programs Division.

Copyright0 1992AmericanInstiNteofAeronauticsmdAstronautics. Inc. NocopynghtisassertedintheUnitedStaresunderTirle 17.U.S. Code. The U.S. Governrnenc has a royalty-free license 10 exercise dl rightsunderthecopy~ghtclaimedhereinforGovernmental purposes. All other rights are reserved by the copyright owner.

g LOX LH2 LRB LV OMS PIA RCS ssv TI-AI

acceleration due to gravity (32.2. Wsec2) liquid oxygen liquid hydrogen liquid rocket booster launch vehicle orbital maneuvering system propulsiodavionics reaction control system single stage vehicle titanium aluminide

INTRODUCTION

As long as this nation, and indeed the world, wishes to maintain a human presence in space, there will be a demand for a system to transport people to and from earth orbit. The goal of

system capable of meeting future mission requirements to m s - pon personnel and payloads requiring a manned presence to and from earth orbit, with increased emphasis on vehicle reli- ability, penonnel safety, and large operational margins. This must be accomplished within the Framework of existing budgetary constraints which are expected to continue well into the next century. Thus. minimal life cycle costs (the sum of design. de- velopment, testing and evaluation (DDT&E), production, and operations costs over the lifetime of the vehicle) become an important design goal.

this nation must, therefore, be to develop a space transportation V

Historically, the competition between the U.S. and the U.S.S.R. has driven both nations to perform manned missions in space. It is hoped now that cooperation with the U.S.S.R.. as well as the rest of the world, will enable man to extend his reach toward the stars even in this era of cost cutting measures around the globe. In the past, military aircraft and space launch vehi- cles have typically been designed to achieve performance and mission flexibility at the expense of operational efficiency. Their designs have been performance driven. Everything from weap- ons delivery systems. reconnaissance aircraft, to expendable launchers, and upper stages have been developed according to this philosophy. The partially reusable Space Shuttle booster/

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orbiter system represents a first step toward designing with the goal of improved operability and reduced life cycle costs. Un- fortunately, current costs for payload to and from orbit using the Shuttle are still too high to make this a commercially viable system. It is clearly evident that in designing the next genera- tion system these factors must be considered from the initial stages of the design process. Routine. affordable access to space necessitates this adjustment from a performance driven design approach toward one driven by operability. This can be likened to the difference between designing a fighter aircraft and com- mercial jet transport. Although the sheer magnitude of the ener- gy necessary to transport material into space makes i t unlikely that we will achieve an airline-like operation in the foreseeable future, a move in that direction is essential if we are to continue mankind’s expansion out into the universe.

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DESIGN ISSUES

Choosing the optimum systedarchtecture to replace the current Space Transportation System (STS) involves several de- sign issues, some of which are technical in nature. but others are related more to design philosophy and matching that philoso- phy to the spirit of the time in which we live. A number of different vehiclescan generally be designed to perform acertain mission, but other factors may ultimately determine what path this nation will follow in the future. Studies‘~’performed over the last several years have served to illustrate the importance of some of these factors, such as mission requirements, design philosophy (operability vs. performance), technology levels, and lifecyclecosts. in theselectionoffuturemannedconcepts. These factors are discussed in a qualitative manner in the next four sections of the paper. Figure I, showing a typical advanced ve- hicle concept, illustrates several related improvements which could be incorporated info any new launch vehicle. More de- tailed trades will be discussed later in the paper where several potential vehicle concepts for the future are discussed.

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Mission Requirements

The selection of “the best” space transportation system/ architecture is highly dependent on the mission model. Too of- ten the mission model is a “moving target.” This has been dra- matically demonstrated with the continually changing Space Station design.’ The orbital altitude and inclination. as well as the size of the station and its components, have been adjusted several times since the inception of the program. Numerous options have been proposed to provide an assured crew return capability. These range from developing a new vehicle, the HL- 20 lifting body. to enhancement of an old vehicle system, Apol- lo shape, and even to the use of an off-the-shelf vehicle, the Russian Soyuz. The Space Station orbital inclination and alti- tude, together with the response- and on-orbit stay-time require- ments and g-force limitations wouldlikely dictate whichofthese concepts would be viable. A space transportation system and architecture based on the current mission model would differ from one that would have best matched the mission model of the original station design.

Mission requirements include the amount of payload for delivery to and return from orbit, the orbital inclination and al- titude, on-orbit stay time, response time, and required crew. Future space transportation systems must accommodate the to- tal payload mass required as well as capture the expected pay- loads in terms of weight and volume. Options which trade increased flight rate for reduced payload capability must be con- sidered. The definition of the target orbital altitude and inclina- tion is also critical to the selection of a vehicle concept, Generally, the payload delivery capability goes down with increase in in- clination and altitude of the required orbit. Missions to low earth orbit (LEO) do not usually require expendable upper stages, but delivery of satellites to geosynchronous orbit (GEO) and inter- planetary missions often do. If these types of missions com- prise a significant portion of the launch requirements a reusable upper stage might be a necessaq part of the space transpom- tion architecture.

Two basic. yet very different types of missions generally must be captured. The first is a rapid-response, orbit-ondemand type ofmission. The second is a regularly scheduled supplyiser- vicing mission which can include boosting of the components for the building of large space smctures. The former requires delivery of a payload to a certain location (orbital inclination, altitude) within ashortperiodoftime.This is typicalofmilitary sorties and crew emergency rescue missions. All-weather capa- bility is critical in order to meet this mission. Also, vehicle turn-

Figure 1. Appropriare mission needs. lessons learned. around time must be minimal. As turnaround time increases a larger stable of vehicles would be necessary in order to satisfy and exploiration of new technologies.

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this requirement. The rapid-response requirement also drives the vehicledesign toward one with extensive loiter capability or the ability to take off and land almost anywhere. Generally, for this type of mission minimal payload capability is necessw. The cost goal for this mission category is, therefore, low dollars per flight. On the other hand. the payload delivery and resupply mission, tends to favor a vehicle design with the capability to deliver larger payloads. Low dollars per pound of payload becomes 3 design goal. All-weather capability and rapid turnaround time. although still desirable, become less critical.

The above discussion indicates qualitatively the importance of the mission model to the choice of a system. The studies in Ref. 1-7 include vehicles designed to boost payloads weighing from 5 to 150 klb. into orbits ranging from 28.5” inclination to polar orbit, at altitudes of 100 nmi to geosynchronous orbit. These studies also address rapid-response, orbit-on-demand missions as well as regularly scheduled supplylsemicins mis- sions. The results often favor different types of concepts de- pending on these mission-related factors. When the time comes for a decision to be made regarding the next generation vehicle, the mission model must be fixed in order to allow the best deci- sion possible. Until that time, vehicle sensitivities to mission model should not be ignored. Depending upon the mission model, entirely different concepts might be favored. Care must be taken not to choose a space transportation system without a thorough understanding of the impact of the mission model on that choice.

Design Approach

In a “design-for-operations” approach the effects of vehi- cle design on recumng costs are taken into account from the outset of the design process. Although higher level studies can be used to gain some understanding of the benefits and features of competing concepts. in-depth analyses are necessq in or- der to adequately understand the technical issues, manufactur- ing and operational complexities, technology requitements. and life cycle costs of promising concepts. Before a concept selec- tion can be made, ground- and flight-operations requirements and approaches, development and pmduction approaches, and mission reliability and safety issues must be thoroughly undet- stood. This understanding and a clear definition of the enabling technologies with the attendant costs and benefits can only be gained through detailed analyses. Life cycle costs will likely be a primary discriminator for architecture and concept selection and the degree of operability and reliability of the system will have a profound impact on those costs.

Experience with the current Space Shuttle system has pro- vidcd many valuable lessons with regard to the factors which must be considered early in the design phases in order to devel- op a vehicle which is operationally efficient. The goals are to drastically reduce the turnaround time required between flights and to improve the safety and reliability of the vehicle so that vehicle replacement is not necessary. These factors can all be improved significantly relative to the current Shuttle system capabilities and must be assessed in terms of impact on vehicle weight and life cycle costs. The degree of improvement possi- ble depends in part on the state of technology assumed. The amount of technology development necessary can greatly im- pact the life cycle costs.

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Figure 2 illustrates a two-stage vehicle “designed for oper- ations.” A future concept designed for operability would have the following characteristics. It would be designed so that the need for between flight inspection. maintenance and repair is minimal. It wouldbeamorerobustvehicle whichdoes not need to be recertified between flights. Use of a durable thermal pro- tection system (TPS) and vehicle health monitoring systems would contribute to achieving this goal. Built-in test equipment (BITE) for monitoring systems would also contribute to the vehicle’s reliability and safety. The potential for parallel pro- cessing of the orbiter and payload would be maximized. Utili- zation of a containerized payload concept would simplify payload integration. Elimination of toxic hypergolic propellants would allow vehicle processing to begin as swn as the vehicle is returned to the launch site. Asingle-stage concept would elim- inate the operations currently associated with mating the vehi- cles. For a multi-staged vehicle, weight of the individual components would be reduced to simplify the mating opera- tions. Staging at or below Mach 3 would eliminate the weight penalties associated with staging at higher Mach numbers. in- cluding the added propulsion to return the vehicle to the launch site and the thermal protection required. Tne vehicle would be

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Mach 3 Slaglng -.. Crew emergency

erape systems

Figure 2. Design-for-operations approach for AMLS W

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either light enough in weight that it can easily be ferfied from landing to launch site or would incorporate a self-ferry capabil- ity. With regard to the safety issue i t would have a single-en- gine-out capability and could either execute a return to launch site (RTLS) or an abon to orbit maneuver. In the event that safe abonofthe vehicleisnotpossibleitwouldprovideacrewemer- gency escape capability.

Technology Levels

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An assessment of the relative merits of competing space transportation systemdarchitectures clearly depends on the tech- nology levels assumed for each vehicle. Typical system studies assume one of three technology levels; current, near-term, or advanced technology The fust includes mature, state-of-the art technolo$es which have been applied to the Shuttle or other systems. They are available “off-the-shelP‘ and require no de- velopment. Near-term technologies are ones which have been demonstrated sufficiently and need little additional development in order to be applied directly to a vehicle within a couple of years. Application of these technologies would require flight certification. Advanced technology is by far the fuzziest term. These technologies require a much more intensive level of fund- ing and study before application to a new vehicle would be fea- sible. Care must be exercised to realistically assess the technol- ogy levels when comparisons are made among advanced vehicle concepts. The concept that appears to be the most favorable at an advanced technology level may not even be feasible at a lower technology level. The assumption of advanced technolo- gy necessarily requires that some estimate be made of the cost to develop that technology and the risks associated with failure to meet the assumed goals.

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Table I lists the three technology levels and the associated structures and materials, propulsion and subsystems assump- tions being used in a current study. 6 State-of-the-art technology

on the Shuttle includes aluminum shuctures and tanks, limited use of composites, reinforced carbon-carbon nosecap and lead- ing edges, ceramiciblanket acreage TPS. space shuttle main en- gines ( S S M E s ) , hydraulic power. hypergolic OMSRCS, and fuel cells.

Near-tern technologies include GraphiteEpoxy compos- ite structures, integral reusable AI-Li cryogenic tanks, titanium or advanced carbon-carbon nosecap and leading edges, durable metallic or ceramic TPS. light-weight L H Z O X fueled SSME derivative. LH2 fueled turhojeflramjet airbreather, and air-tur- borccket propulsion. All-electric systems using electromechan- ical actuaton replace hydraulics for engine gimbal as well as aero-surface controls. Hypergolic propellants are replaced by v

Table i. Technologiesfor Eanh-to-orbir vehicle options.

cryogenic hydrogen and oxygen in OMS, RCS and fuel cells. Advanced avionics and fault-tolerant, self-check subsystem components for vehicle health monitoring are included at this technology level.

Complementing the near-tern technologies identified in the previous paragraph. advanced technology assumptions in- clude TI-AI composite smctures, insulated in some areas to act as a TPS as well, non-integral reusable thermoplastic hydrogen tanks, variable mixture ratio rocket engines using slush hydro- gen and triple point oxygen propellants, and ramjekramjet propulsion. Vehicle subsystems are similar, but lighter in weight due to the use of advanced materials. Actively cooled inlets and nozzles. and semi-active heatpipe systems are also assumed at this technology level. Some technologies such as reusable cryo- genic tanks are seen to be enabling. Le. essential to the design and development of a new system. whereas technologies such as electromechanical actuators are enhancing, i.e. they improve the system either by reducing weight or improving operations and reliability, but failure to develop such technologies would not preclude the development of a new vehicle.

The concepts assumed for the advanced technology level require “technology investment” now in order to be available for the next generation manned launch system. The ALS (Ad- vanced Launch System), NLS (New Launch System) and the NASP (National Aerospace Plane) programs have each con- tributed to the funding of advanced technology programs at some level. Some of the technologies discussed are applicable to any future system and should be developed regardless of the con- cept to be chosen for the next manned launch system. These common technologies include, reusable cryogenic propellant tankage, operablelreusable main propulsion, vehicle health

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Cost' Considerations

U Two major motivating factors are forcing this nation to con- sider replacement options for the current Shuttle fleet. One is that the fleet is aging. Considering the long lead times needed to develop the technology and bring a new system on line. it is critical that this effort begin now. The other relates quite simply to economics. The Shuttle with its labor-intensive maintenance, repair and certification requirements, does not provide the com- mercial viability hoped for at the inception of the program more than 20 years ago. It is simply too expensive to operate. Numer- ous technological advancements over the last two decades now offer the potential to design and build a new system with drasti- cally improved operational characteristics. The hope is that the costs over the system's life cycle from development through the operational years can be significantly improved when com- pared to those for the Space Shuttle and expendable launch ve- hicle (ELV) systems currently in use. Some have suggested that an order of magnitude reduction in launch costs may be achiev- able. Figure 4a from Ref. I O illustrates a possible breakdown in the ways this might be achieved. At the same time n dramatic reduction in recurring costs is also a goal. Figure 4b shows the

monitoring and management, long-lifeflow-maintenance dura- bleTPS, autonomous flight controUachptable software and avi- onics, and operationally enhanced subsystems. Other technologies are more concepVvehicle dependent and contin- ued funding of the development of these technologies should be closely linked to a clear payoff for their application to a fu- ture system, either in terms of enhanced capabilities allowing the performance of critical missions or in terms of commercial viability. System specific technologies include airbreathing pro- pulsion, active and semi-active cooling, and variable mixture ratio propulsion.

Typically advanced vehicle system studies assume a mar- gins on the order of I O to 15 percent in weight included in the final weight estimate. This margin is intended to account for the apparent weight growth which typically occurs as more detailed analyses are applied and a concept matures. Often advanced technologies are applied by using a percentage reduction in weight, depending on the technology level assumed. For exam- ple in Ref. 9,25- and SO-percent reductions in weight from the SSME engine weight were assumed for near-growth and ad- vanced technology level engines respectively. Historical trends. however, indicate that both the margins and the weight reduc- tion factors assumed for technology improvement may be opti- mistic. Figure 3 illustrates that a margin of 25 percent might be more realistic. Care must be taken in applying a blanket margin value in all vehicle studies. This could have the effect of unjus- tifiably biasing the results in favor of the more advanced tech- nologyfless mature vehicle concepts. The less mature the concepr the greater the margin that should be applied. It is therefore, not unreasonable to apply one margin to a vehicle which utilizes existing or near term technology and a greater margin to one incorporating more advanced technology.

6o r

-101 1 0 20 40 60 80 100

Program completion, percent (contract authority 10 proceed to 1k.t llighl vehicle1

Figure 3. Sumrnav of 18 DoD, commercial, and NATO space vehicle weight growths.

5 w o r Approximate Oirtributioo

01 me 90% Cost Reduction

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Figure 4a. Factors relating to launch COS< reductions

SHUTTLE (STS) SHUTTLE II

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Inlerencs: Full reusabiiify Simple operations

Figure 46. Inferences of launch COS[ reductions on advanced sysrems.

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current breakdown for the Shuttle system compared to the mag- nitude and the break down we would hope to achieve with a new system.

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The goal for a replacement system will be to minimize life cycle costs, retain and improve mission capability, and improve operability, reliability and safety relative to thecunent system. Costs in a space transportation system development and opera- tions progam can be categorized in two ways; DDT&E costs incurred up front. and operations costs over the life of the vehi- cle. The total of the two over the life of the vehicle represents the total life cycle cost. Minimization of the life cycle costs requires that the DDT&E and the operations costs be balanced in an optimal way. Typically, money spent up front on DDT&E will save money in operations costs later. Conversely, money saved up front in DDT&E costs will often increase the operations costs over the lifetime of the vehicle. Even after program completion it is difficult to determine in retrospect what the optimal split should have been. Attempting to do so during the preliminary design phase is a monumental, if not impossible task. Only in the last several years have efforts been made to include some kind ofrealistic cost estimates in the preliminary design studies. With theenormouscomplexity oftheproblemofpredictingcosts, the best we can currently hope to do is to identify the cost drivers and use engineering judgment to design the vehicle system and architecture to minimize life cycle costs. v

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Figure 5. NASA rechnology reodiness levels.

creases the DDT&E costs even more. The degree to which this nation is willing to support advanced technology development may vary weli 6e the factor that determines what the next gen- eration of manned launch system will look like.

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Vehicle complexity also impacts DDT&E costs. The more complex the vehicle and its subsystems, and the more compo- nents in the system, the greater the integration difficulties and the higher the DDT&E costs. Concepts which require the de- velopment of multiple parallel systems also tend to have higher DDT&E costs (e.g. design and testing of two different engines is much more costly than simply designing and developing two of the same engine). The reliability built into the system in or- der to eliminate the need for repair, replacement or recertifica- tion between flights and the testing strategy required to verify that reliability are costly in terms of DDT&E. If these issues are addressed during the design and development phase, operations costs may be reduced over the lifetime of the vehicle. Figure 6 illustrates how life cycle costs tend to decrease with increased reliability until a level of about 98.5 percent is reached. This decrease is due primarily to reduced operations costs. Beyond 98.5 percent the increase in DDT&E and production costs to improve reliability result in a significant rise in life cycle costs. The degree of reusability also impacts the DDT&E costs be- cause it is more difficult to design and cenify a reusable vehi- cle. Operationally though, this means fewerexpendable vehicle components which must be replaced each mission and lower recurring costs. Finally, the vehicle size plays a role in DDT&E costs. Larger facilities to handle a larger vehicle cost more to build, operate and maintain. However, at a given technology level increases in dry weight are a weak measure of increased DDT&E costs. In general though, the technology level is the dominant meuic. Often a low dry weight is achieved at the ex- pense of high technology development costs required to devel- op the very systems which have reduced the weight.

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4 cost

.95 .96 .97 .98 .99 1.00 Vehicle reliabilily

Figure 6. Cosrs of vehicle reliability.

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The program management approach utilized in terms of development schedules and programmatics can also play a ma- jor role in determining the DDT&E costs. Arecent study'' has compared the use of a Skunk Works (Sw) approach, as used by the Lockheed Advanced Development Company, with that of a Business as Usual (BALI) approach for development of the HL- 20 Personnel Launch System (PLS). The SW approach pro- vides foremphasisonearlytlight testingversusextensiveground testing and maximizes the use of "off-the-shelf' equipment wherever appropriate. This represents a higher risk during the early pan of the development phase than the BAU approach, but a lower risk and more mature system as thz system evolves to become operational. This approach provides opportunity for program cancellation at critical milestones prior to full commit- ment of development funds. The BAU approach, used in the Shuttle program, entails much more ground testing before flight of a protorype vehicle as well as commiment of a much greater portion of the development funds before an fully integrated sys- tem is demonstrated. These approaches are described in more detail in Ref, 11. Figure 7 shows a comparison of the cumula- tive DDT&E funding for the two approaches. For the HL-20 vehicle this study showed that the SW approach shonened the development schedule by 21 percent and reduced the develop- ment cost by 44 percent over the BAU approach. thus indicat- ing that programmatics play an important role in development schedules and costs. It should be noted however. that the SW approach is best suited to small limited production systems with well-defined requirements, few required organizational inter- faces and where advanced technologies are not a major require- ment for development.

Several of the factors which influence operability and thus operations costs have been discussed previously. The difficulty of the handling of toxic and potentially explosive propellants

First RrSl Orbital Right Orbital night

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F i p r e 7. PLS DDTBLE cumulativefundiny.

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Figure 8. Vehicle concept integration costs

incresses operations costs. Vehcle integration requirements play a l q e r o l e a s well.Themorecomponentsthemoredi~cult the operations and thus the greater the cost. Figure 8 illustrates the reduction in components possible relative to the current Shuttle system. The operations costs for expendable flight components are functions primarily of the flight rate and payload require- ments. The greater number of flights and the greater the amount of payload canied to orbit the higher the operations costs. Op- erations costs are also impacted by the trade between flight rate and vehicle payload capability. A vehicle with sea te r payload capability can deliver the same amount of payload to orbit with a fewer number of flights. Whether this is cheaper operationnl- ly or not would require a detailed definition of the architecture as well as the vehicle concepts. For reusable vehicle compo- nents the operations costs are related to the between flight in- spection, repair and replacement requirements. If these factors can be minimized through design of a very robust vehicle which only needs recertification after a major in-flight anomaly, oper- ations costs will decrease dramatically. As mentioned previous- ly, the DDT&E costs required do, however. increase with the reliability or robustness of the system. The manpower require- ments for nomal servicing between flights represents an area where significant operations cost reduction can be achieved rel- ative to the Space Shuttle system. Normal servicing require- ments between flights include standard turnaround time. The greater the turnaround time the greater the operations costs. Turnaround time represents downtime during which the vehi- cle cannot be performing its mission. Finally, the need to ferry the vehicle back to the launch site represents another operation- al cost. due to the need to mate and demate the two vehicles and to provide a crew to ferry it back.

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7

BASIC SYSTEM CONCEPTS -THE OPTIONS d

Background

Several studies performed over the last decade or so have ex- amined possible concepts forthis nation's next generationmanned launch system. The results of these studies form the basis for the comparisons and trade$ discussed in this section. The fmt of these, the Future Space Transportation System study (FSTS).I was ini- tiated in 1981. the year ofthe firstsuccesshl Shuttle orbital flight. Shown in Fig. 9, the vehicle concept chosen was a tw+stage. re- usable orbiter and booster capable of uansporting 150 klb to earth orbit The Orbit On Demand Vehicle study (OODV)? initiated soon after the completion of the FSTS study, examined a broad range of vehicle options designed to deliver a 5 klb payload to orbit in a rapid response mode. Figure IO illusmtes the various con- cepts studied which include expendable and reusable, one-. hvo-, and even rhree-stage, vertical and horizontal launch concepts.

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$KL Gmrnwetght,i;lb ......... 4.897 oltmr ....................... 2.747 8W1181 ........ ............ 2,150 LH2

I--- [sen -I Figure 9. Configuration design for operations

and performance.

The Shuttle 113 study, initiated in 1985, looked at vehicles to satisfy two different missions: one to deliver large masses (bulk cargo, propellants, large satellites) at low cost per pound and the other to perform rnanndsorlie missions. This study complemented the Space Transportation Architecture Study (STAS) and a variety of options were considered to deliver a 12 klb payload to polar orbit and one of 100-150 klb to LEO. Figure 1 I illustrates the phased approach examined by which capability could be gradually expanded beyond that of the cur- rent Shuttle and expendable fleet. 'W'

propillants

<-+ v.1 * - Vertical-takeoff single-stage concept.

Vertical-takmff two-stage eoncept with drop tank.

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Horizontal-takeoff single-stage eoncent. A sled would be used for takeoff.

- . - = - < - Horizontal-takeoff two-slaw c o n ~ q t with supersonic staging.

Horizontal-takeoff two-stage A C O ~ C C P ~ with hypersonic staging.

Horizontal-takeoff two-stage concept with subsonic sfaging

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Horizontal-takcoff two-mgc :oncept with drop tank.

Horizontal-takeoff thrcc.stagc :onccpt.

Figure IO. Orbit-omdemand vehicle conceprs.

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, -. . . LOWLH2 core LOXILH2 core LOWLHP core Snunle I1 snme iI stage + solids + ~oos1er + PLS vehicle 2005 + PLS vehicle 1995-100klb 1998-150mlb 1998 Deliver/ 2005

8 5erVIcing PelSOnneI tianspon

Figure 11. Phased approach rransporfarion archirecrure.

Government baseline configuration

Podded vehicle

Conical vehicle

Figure 12. Early NASP vehicle conceprs.

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Figure 13. Artist's conception of rhe NASP vehicle

The concepts investigated in each of these studies were rocket powered with the exception of those in the OODV study which included airbreathing first stages. The beginning devel- opment ofscramjetenginescapableofhigh Machnumber flight lead to study of asingle-stage airbreathing vehicle, NASP4 Fig- ure 12 illustrates three single-staged. all-airbreathing, horizon- tal takeoff and landing concepts proposed to satisfy the NASP mission to deliver a crew of two to orbit. These have evolved into the configuration shown in Fig. 13 currently under study.

W In 1991 industry was asked to propose possible concepts

for a future manned single-stage-to-orbit rocket-powered ve- hicle for the Strategic Defense Initiatives Office (SDIO). The proposed concepts, shown in Fig. 14. included one horizontal and two vertical landers. The McDonnell Douglas DC-X which recently flew its second low altitude demonstration test in less than a month is a small scale experimental version of the vehicle shown at the right in Fig. 14.

Figure 14. NASA evalutarion of SDlO Phase I SSTO concepts. v

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The studies mentioned here have contributed to an exten- sive knowledge base with respect to the options for future manned space transportation systems. Since 1988 efforts have been focused on an activity known as the Advanced Manned Launch System (AMLS) study6 This activity continues today and represents the culmination of these studies. It is described in the next section and provides the basis of the vehicle concep- tual trades discussed in this paper.

A M J S

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In the 1988 time frame the Shuttle E effort was made part of a larger NASA Headquarters study to define options for the next generation manned space transportation system. The goals of this AMLS effort are to define systems that meet required mission needs in terms of personnel transport and manned-pres- ence-required payloads, improved cost effectiveness, increased vehicle reliability, and large operational margins. Three basic

Primary RCS

Figure 15. HL-20 PLS vehicle concept.

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*Ir i""'" TwOI"" s,"gIe-stage

c/ Figure 16 Manned concept options

approaches are under consideration; I ) evolution of the current Shuttle through subsystem and block changes, 2) definition of a small PLS for canying people and small amounts of cargo. and 3) development of an all new vehicle. The HL-20 concept' shown in Fig. 15 has been studied in great depth as one option for the PLS mission. Figure 16 illustrates the broad range of concepts being examined in the AiMLS study. They include ex- pendable and reusable, one- and two-stage, rocket and airbreath- ing concepts. References 5 , 13. and 14 describe in detail the concepts considered and the trades that have been performed.

Mission: ConsistentwithNASArequirements andotheron- going studies, two mission scenmos have been assumed for the vehicle conceptual studies. The first is a due-east space station logistics mission from Kennedy Space Center (KSC) and the second apolarpayload delivery mission from Vandenberg West- em Test Range (WTR). Table 2 describes the design parameters for each of the missions. A crew of 10 was baselined for the logistics mission and all concepts were required to have crew- escapecapability, singleengine-out capability for each stage. and a 15 percent dry weight growth margin. A crew of 2 was base- lined for the polar mission and no crew-escape or engine-out capability was required. The dry weight growth margin was re- duced to 10 percent. The 10 klb payload requirement to polar orbitvanslatedto30klbtoandh.omthespacestation.Therefore, the vehicles were also required to land with a 30 klb payload.

Table 2. AMLS space station logistics ondpolar payload deliveq mission deBnitions

Mission

Payload, Ib Orbit. nmi inclination, deg Mission duration, days Payload dimensions, ft On-orbit AV, Wsec Launch site Landing site

Space Station Logistics

40.000

220

28.5 3

15x30

1350

KSC KSC

Polar Payioad

10,000

1 2 x 2 0

850 WTR WTR

Technology Levels: Two technology levels have been se- lected to investigate their impact on the various AMLS con- cepts. These two levels are designated "near term" and "ad- vanced" and, together with the materials, propulsion and subsystems associated with each. were described in some detail in the previous section on "Design Issues" (see Table I). The identification and quantification of the potential benefits of new technologies is a primary focus of the AMLS study.

10

Design Concepts: Early studies addressed the expendable versus reusable trade. Figure 17 presents the perceived wends in Life cycle costs with varjing degrees of reusability. The intercept represents the DDT&E costs plus the cost of the fmt production units and the slope represents the recurring costs. It is apparent that for limited launches over the life cycle, expendable systems are desirable. However, as the total number of launches increases the reusable systems may become more cost effective. .All of the vehicle concepts presented here will therefore be fully reusable. Figure I6 presents the range of vehicles which have been consid- ered in the AMLS study. Winged configurations have ken con- sidered for all vehicle elements because of the operational benefits gained from low e n e j g-load conditions, high crossranee. hori- zontal landing, and horizontal takeoff (for airbreathmg concepts). l7 ,.///-. Expendable ./ >zb

costs Fuily reusable

Total launches over lifeqcle

Lifecycle _/--

&--

Figure 17. Effects of vehicle reusability on &cycle cost rends.

Concepts for the Space Station Logistics Mission: Sin- gle- and two-stage vertical takeoff, winged rocket concepts were considered for near-tern and advanced technology levels. Air- breather systems were not considered for this mission. Table 3 summarizes the characteristics of the vehicles sized for this mis- sion at both the near-term and the advanced technology levels.

Table 3. AMLS options for space station mission.

G- weight. Klb Ory weqht. KJb Baav lengm, fl

~%%wQhlUlb Ory weight, Kib I 1:

%dy length. fl PmwlSlcm me D = 551150

'SSME dewatwe

'SSME weigh15 reduced by 25%. 1: "SSME weights reduwd by 50%. 1

Tunnel & 15-11 X 304 Jettisonable ybf, payload bay crew moduie 7

LHZIank-.

GLOW: 2552 Klb Dry wt: 336 Klb

, I I I IL-

U

Elevons

Eody flaps

* Vertical takeoff * LOWLHZ SSME-derivative engines

* Parallel bum wilh crossfeed * External payload canister with aerodynamic shroud

Figure 18a. Near-term technology No-srage MLS

Unmanned glideback booster: Mach 3 staging

(space starion mission). v The near-term technology two-stage vehicle concept is il-

lustrated in Fig. 18a. The LOX/LH2 fueled SSME derivative engine is represented by a 25 percent dry weight reduction with a 15 percent margin. The vehicle's gross weight is a little more than half that of the Shuttle system (2.6 Mlb vs. 4.5 Mlb). Its dty weight is 340 klb. The design issues for this type of config- uration include engine gimbal requirements prior to staging, the staging maneuver and the aerodynamic effects of the mated configuration.

The single-stage vehicle concept at the near-tern technol- ogy level is an adaptation of the orbiter stage from the two- stage system. The vehicle designed to meet this mission with safety. reliability. and operability built in is seen to be totally infeasibleat this technology level (seeTable3). Its gross weight of25 Mlh, dry weight of 2.8 Mlb and length of 390 ft dwarf the other concepts. However, single stage to orbit (SSTO) vehicle sizing for these conditions is extremely sensitive to small weight or mass ratio changes. Aone-percent decrease in mass ratio cut the weights shown in half. This sensitivity to weight growth would also translate into high risks in development, production and costs. Development of an SSTO of this configuration for [his mission using near-term technology would be risky at best. -

The advanced technology two-stage concept is shown in Fig. l8b. The slush LHZtriple-point-oxygen fueled S S M E de- rivative engine is represented by a 50 percent dry weight reduc- tion with a 15 percent margin. With a gross weight of 1.5 Mlb and a dry weight of 180 klb this vehicle is considerably lighter than the two-stage vehicle at a near -term technology level.

W

Slush hydrogen .&oxygen lams

GLOW 1543Klb D y Y : 178Klb

Triple-point oxygen lank

Elevon

Body flap

* Verticai takeoff Slush LH2; SSME-derivative engines Unmanned giideback booster: Mach 3 staging - Parallel bum with crossfeed

0 lntemai payload canister 'd

Fi,yire 18b. Advanced iechno/o~)' rwo-srage .4MLS (space station mission).

v

Figure 18c illusuates the AMLS SSTO concept when ad- vanced technologies are applied to the near term system. The advanced propulsion concept is represented in the same way as for the advanced technology two-stage design. Although with a gross weight of 2.7 Mlb and a dry weight of 250 klb. it is still significantly larger than the two-stage system at the same tech- nolog level. It is, however, significantly lighter in weight than the near-term SSTO and in fact its dry weight is less than that of

179 fl -4 GLOW: 2654 Klb Drywt: 251 Klb Elevons

Body flap

Vertical takeoff . Slush LH2; SSME-derivative engines f External payload canister with aerodynamic shroud

Figlire 18c. Advanced iechnolo&y single-srage ..IMLs (space station mission).

Table 4. AMLS options for polar mission

465 357 1359 1334 62 57 125 112 122 138 144 135

c.551 &=55 I e-551 c - M

'SSME "SSME "SSME VMR l e m s demw8 d e w w e

150 150 rsa 120

12

the near-ten two-stage vehicle. Considering the advantages in operations for a single component system, this would likely be a competitive system provided the technologies are cost effec- tive to develop.

Concepts for the Polar Payload Delivery Mission: The near-term and advanced single- and two-stage rocket vehicles were resized for the polar mission. In addition, four sirbreath- ing concepts were also sized. These concepts include a two- stage turbojet/ramjer/rocket-pmpelled system at near-term and advanced technology, and low-speed cycle/rarnjet/scranijedrock- et SSTO vehicles at an advanced technology level. Tahle 4 from Ref. 9 summarizes the characteristics of the vehicles sized for this mission at both the near-term and the advanced technolog) levels. For this mission the weight reduction for the S S b E de. rivative engine was assumed to be 25 percent and 50 percenl respectively for near-term and advanced technology. A IO per- cent margin was assumed in both cases. Figures 19a and b illus- trates the comparative sizes of the different vehicle concepts foi near-term and advanced technology respectively.

Tweslags rocket

Tweslage airbrealhedrockel

SSTO m k e l kd=b

Dry weight. klb

167

440

427

I 0 100 200

Length. H

Figure 190. Near-term technology AMI3 concepts (IO klb polar mission) to same scale.

ATWIocke1 [ 214 ssv I - ,, 157 conlca A0 ssv I

0 100 2 w 300 Length, fi

Figure 196. Advanced technology AMLS concepts (10 klb polar mission) to same scale.

For the polar mission the rocket vehicle systems are rough- ly halved in both gross and dry weight. The conclusions reached for [his mission are generally the same as for the logistic mis- sion except the near term SSTO has decreased in weight more dramatically than the other vehicles due to the extreme sensitiv- ity mentioned previously. Various factors influencing the rock- et vehicle sizing are illustrated in Fig. 20. The vehicles designed for the polar mission are designed for performance and the weizhts would grow accordingly if the operations characteris- tics required for the logistic mission were factored in. It is also apparent that the SSTO has become more competitive with the utilization of a design for performance philosophy. In fact the advanced technology SSTO concept which utilizes a variable mixture ratio (W) SSME derivative is weight competitive with the two-stage system and would likely have the advantage in operations.

w

f Vehicle gross 'wan weight

v Advancing technology --f

Design lor performance --t 4- Design for operalions. safety, reliabilily f Increasing payload. margins

Figure 20. Rocket vehicle sizing considerations

Single- and two-stage. near-term and advanced technolo- gy airbreathers are considered for the polar mission. All air- breather concepts are designed to takeoff and land horizontally. The near-tern two-stage concept employs turbojets to Mach 3 and ramjets until smging at Mach 6 where the orbiter is pow- ered by SSME derivative rockets. Table 4 shows that the gross weight of the two-stage airbreather is less than that of either the single- or two-stage rocket system. The reason for this is be- cause the oxygen needed is obtained from the atmosphere rath- x than canied on board. The dry weight, however, is higher than either of the near term rocket systems. This is due primar- ly to the increased weight of the airbreathing engines and the drag losses associated with this type of concept. The advanced two-stage airbreather concept is similar to the near-term one, except that advanced technologies are employed. As was Vue o i the near-tern system, the gross weight of the advanced two- stage concept is less than either rocket system. but the dry weight

v

13

is greater. Thus the two-stage airbreathers do not show a dry weight advantage over the rocket system at the equivalent tech- nology level. In fact, it can be seen from Table 4 that the two- stage airbreather is actually higher in dry weight than the either of the SSTO airbreathers. This result is atuibuted at least par- tially to the drag losses. The requirement for RTLS with the orbiter attached increases the wing design loads and the sttuc- tural weight in turn.

d

Two advanced technology SSTO airbreather concepts were also considered. The air-turborocket (ATR) concept is shown in Fig, 21a. Similar to the British HOTOL. this vehicle utilizes an unpowered trolley during the takeoff roll. The ATR powers the vehicle to Mach 6 where a rocket engine takes over. This vehi- cle is described in greater detail in Ref. 16. Once again. it can be seen from Table 4 that the gross weight is less than either rocket SSTO, but the dry weight is significantly greater even though the uolley weight has not been included. The other advanced technology airbreather SSTO is illustrated in Fig. 2 I b. It is a conical accelerator vehicle which utilizes a multi-mode slush hydrogen-fueled propulsion with low-speed, ramjet, scramjet, and rocket cycles. The vehicle forebody is a 5O half-angle cone that compresses the flow before it reaches the inlet. This vehi- cle is described in greater detail in Ref. 4. Table 4 indicates that this vehicle has the lowest gross weight of all the SSTOs con- sidered. It is slightly lower in dry weight than the ATR concept, but the dry weight is nonetheless greater than the advanced tech- nology rocket systems.

'4

Elevon GLOW: 1087Klb Dry w: 214 Klb

Tip fin OnlrOlle,

'd

Mullirarnp inlets J' ATR engine

Horizontal takeoff using unpowered trolley

engine from Mach 6 to orbit - Slush H2: air-turborccket engine to Mach 6: racket

. Internal payload canister with aerodynamic shroud

Figure 21a. Advanced technology air-turborocket'rocket SSTO (polar mission).

,--== 3 crew ,

n _ , - D GLOW 451 Klb Drywt: 157Klb

,- Slush hydrogen lank Payload /

/

Mulli-cycle engines (360' wraparound) I

- Horizontal takeon * Slush HZ

Multiple wraparound engines Internal payload canister

Figure 21b. Advanced technology airbreathing conical accelerator SSTO (polar missionj.

Single- vs. TwoStage Concept Analysis

The studies described here provide some answers to the question of whether future systems should have multiple stages or not. Clearly, SSTO rocket systems of the configuration as- sumed for large payload missions are not feasible assuming near- term technology levels. The size and weight disparity between one- and two-stage concepts increase with increasing payload and crew size as well as with mission requirements biased to- ward operability. If a system were to be built with near-term technology, a two-stage rocket system would be the only real option. The choice becomes less clear with the evolution of ad- vanced technology. For high flight rates and advanced technol- ogy,theoperationsadvantagesofeitherarocketoranairbreather SSTO may outweigh the dry weight penalty relative to a two- stage rOcket system. For low flight rates the decision is less obvious and a detailed analysis of the costs is necessary.

The two-stage airbreather is shown to be higher in dry weight than the SSTO airbreather. Numerous factors including drag losses associated with the complicated aerodynamic de- sign, RTLS requirements, increased heating, and subsystem du- plication on the two stages combine to cancel out the advantages of staging seen in the rocket systems.

Rocket vs. Airbreathing

The choice between rocket systems and airbreathing sys- tems is equally complex as evidenced by the results obtained in the studies described here. The nex-term two-stage and the ad-

14

vanced single- and two-stage concepts clearly show il gross weight advantage over the rocket systems. This does not trans- late into a reduction in propellant costs as one would hope. Drag losses in the airbreathers force the vehicles to carry more pro- pellant. LH2 is much more expensive than LOX and resuk in the propellant costs for the airbreathers actually beins greater than for the rocket systems. The explosive potential oi the air- breathers is, however, somewhat less than that fora rocket sys- tem without the close proximity of the oxidizer and the fuel. This gives the airbreathers a slight safety advantage. Without exception, the dry weight of the all-rocket vehicles is less than that of their airbreathing countelpars. The complexit]i of the airbreathing propulsion systems relative to the rockets would likely make their DDT&E costs greater. In terms of mission related capabilities, the rocket systems are best suited to civil- ian missions requiring payload delivery to orbit. Military mis- sions that stress rapid rendezvous and orbital intercept. loiter. recall, or self-ferry capabilities favor airbreathing systems.

Design for Performance vs. Design for Operations

An assessment of the impact of design approach on he near term SSTO mke t vehicle b m Table 4 was performed. Figure 2 2 1 illustnres the gross- and *-weight reductions t o m he original concept as advanced technologies are included to enhance h e per- formance. The relative effcts of each technology are deceiving becausethevehiclesensitivity isdecdwiththeadditionofeach technology application. The vehicle dry weight is reduced from 430 klb to 70 klb in his exercise. This represents an experimental vehicle design philosophy where each technology is applied to re duce vehicle weighr Fig. 22b illushates the effect of uppding the 70 klb vehicle in Fig. 22a to improve its operability. After all of the operability factors have been applied, the final vehicle d v weight is 130 klb, 3cO klb less than the original near term vehicle concept

5000 r 4000

G m r a 3000

X l b 2000

1000

0

weqn,.

2WO r

Figure 226. Design for operations rocker SSTO vehicle.

Launch Mode Options

Two basic launch mode options exist; ground and air launch. The ground-launch options are either horizontal or vertical take- off. Historically, horizontal takeoff has been the mode for air- craft operations. with helicopters and short-takeoff-and-landing vehicles being the exception. On the other hand, vertical take- o f f has been the preferred mode fur rocket launch systems. The primary advantage of the vertical-takeoff, ground-launch mode is that the landing gear and wings do not have to be sized to support the vehicle at its gross liftoff weight. This results in significant dry weight savings. The disadvantage is that very expensive launch facilities have to be constructed, maintained, and operated and situated so that the desired orbital inclinations can be reached with a reasonable payload and without requiring overflight of inhabited land masses. The vehicle must be able to either return directly to that launch site or be ferried there.

L/

The air-launch mode could offer several advantages. In terms of mission. it would allow an offset launch, loiterirecall, self ferry capability, have basing flexibility with minimal envi- ronmental impact.Airlaunch couldbeaccomplished using near- term technology. Staging provides less sensitivity to dry weight growth than SSTO. This launch mode would also offer benefits in performance in terms of reduced gravity and drag losses, improved rocket performance, reduced dynamic pressures on the orbiter, and reduced orbiter size. Operations would be sim- plified and more aircrah like. Air launch would provide flexi- bility in terms of abort as well as offer improved safety and reliability. If the canier aircraft could be derived from an exist- ing aircraft or technologies, development costs would be mini- mal.

Figure 22a. Design for p e r f o m n c e rocket SSTO vehicle.

The feasibility of an air-launch system depends on a num- ber of parameters. These factors include the mission require- ments such as payload mass and size, cariiercruise requirements, and staging Mach number. The degree of vehicle reusability. the vehicle technology level, and systems required (crossfeed. rockets on canier aircraft, manned systems) also impact the fea- sibility. The aircraft carrier weight capacity is critical and the feasibility may depend on the development of this vehicle, be it an existing vehicle which needs modification or an entirely new vehicle. In order for an air-launch system to achieve its full po- tential it must fall within the weight bearing capacity of existing runways.

~d

Options for air-launch. including subsonic, supersonic and hypersonic staging, have been investigated. Subsonic air launch is feasible for reusable systems only with advanced technology and modification of existing aircraft. Supersonic staging would require the development of a new carrier aircraft. Smaller gross weight systems could be achieved, but dry weights would in- crease. Aerodynamic issues including transonic drag, vehicle integration, and separation dynamics and control are complex for supersonic and hypersonic staging. For hypersonic air launch, operations advantages are not apparent. Some mission advan- tages would be gained.

Air launch systems may have a place as future launch sys- tems. but many concerns must be addressed. These include: pay- load limitations, aerodynamic integration, safety/reliability of the swging maneuver. carrier aircraft modificationddevelop- ment. abort scenarios and cost. Operational advantages over rocket- or single-stage vehicles have not been demonstrated for larger payloads or manned missions.

v

Access To Space

In early 1993 the NASAadministrator initiated the Access TO Space (ATS) activity” in response to a request written into the 1993 NASAappropriation bill. This request in part directed NASA to “recommend improvements in transportation and Space Station” and to “examinefrevalidate civilian and defense programmatic requirements.”The goal of the ATS activity is to define a National strategy to meet future space transportation needs, with the main focus on improved reliability and crew safety and definition of engineering approaches that will offer substantial reductions in operations costs. Peopldpayload re- quirements must be met and margins increased relative to the current Shuttle system. Similar to the AMLS activity, the ATS study is focusing on the three approaches shown in Fig. 23. The ultimate goal of this effort is to lower the cost of routine access to space so that this nation can continue to expand the frontiers d

* Satisfy peapielpayload requirements

e Improve ms1 effeniveness

Increase rel,abfllly

*Increase margins

ADVANCED MANNED LAUNCH

SYSTEM - Clean Sheet STS

STS EVOLUTION

-Evolve existing system

PERSONNEL replacement LAUNCH SYSTEM

*Separate pwpie from cargo - Complement STS

Figure 23. Possible scenarios for the nexf manned space rransponarion syrem.

of space. The system studies described in this paper provide much of the background for the ATS study.

Concluding Remarks

Ultimately, the question is not whether this nation will make that next step into space, but when and how. We are now poised at acrossroads in our nation’s history. The Shuttle fleet is aging and if we hope to maintain an active presence in space the deci- sions must be made now concerning the road we shall take. That decision will be an incredibly complex one. In the early days of the space program the answers were simple. A single mission was to be performed and the state of technology w a such that mission could only be accomplished one way. The nation’s budget and enthusiasm did not realistically limit the way in which a mission could be performed.. In short, the de- signs were driven by performance and not limited by budget constraints. Today things are different. There are more choices, but there are also more constraints. Mission requirements, de- sign approach, technology level assumptions, and cost consid- erations will all enter into the determination of what the next manned launch system should be. Each of these issues may drive the design in a different direction. The challenge is to find the optimal trade among these issues.

Minimization of life cycle costs appears to be the metric which will be used to determine the concept most suitable for the next generation space transportation system. Minimum life cycle costs will be achieved through the optimal balance of DDT&E and operations costs. To “design for operations”, yet satisfy the mission requirements, rather than to simply ”design for performance”. appears to be an imponant key to reducing

16

life cycle costs. Numerous studies performed over the last de- cade have investigated a broad range of vehicle concepts and missions at different technology levels. These studies, cspecial- ly the AMLS study, have provided the experience necessary to make an educated decision with regard to the next generation system. Issues such as expendable vs. reusable systems. single- vs. two-stage systems, rocket vs. airbreather, horizontal vs. ver- tical takeoff, etc. have been addressed and all within the context of assumed technology levels.

Investment in technology development now is critical if this nation hopes to have a cost-effective, operable, and reliable vehicle capable of meeting its launch needs in the early p a t of the next century when the aging Shuttle fleet will necd to be replaced. This technology development has begun under the NASP, ALS. and NLS programs. The studies described here show [hat either a two-stage rocket system, or an architecture based on expendables complemented by the development of a PLS and a Cargo Logistics Vehicle, would be the choice at cur- rent technology levels. Either system would represent an im- provement in operations over the Shuttle system, but not to the degree that could be obtained with some investment in advanced technology. At an advanced technology level SSTO rocket sys- tems and to a lesserextent SSTO airbreathers can become com- petitive. If a vehicle were be to designed for performance alone the airbreathing concepts might have an edge. However, kom a life cycle cost point of view, the rocket system would be the likely choice based on the lower level of complexity and more mature technologies.

This paper has explored the design issues critical to the selection of the nation’s next manned @ansportation system. Sys- tems studies performed over the last 15 years have enabled us to gain much better understanding of the trades involved. The Access to Space effon currently under way is examining a cur- rent-, a near-tern and an advanced technology option for a fw ture space transportation architecture. Three concepts are being considered for the advanced technology option. These concepts will be used to assess the technology benefits with respect to operability, reliability, safety and costs for the next generation system. Perhaps the Access to Space effon will be the one that opens the next door to our future.

Acknowledgements

The work presented in this paper represents the effom of several members of the Vehicle Analysis Branch at NASA Langley Research Center. The authors would like to recognize the contributions made by Chris CNZ, Walt Engelund. Roger Lepsch. Jim Martin. Mark McMillan. Doug Morris. Chris

Naftel. Dick Powell, Doug Stanley, Howard Stone, Ted Talay, George Ware, and Nancy White. Other contributors include Anne Costa of Mason & Hanger Services, Inc., Arlene Moore of Lockheed Engineering and Sciences Company, Garry Qualls of Wgyan, Inc., and Jim Robinson of Old Dominion University.

I

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3.

4.

5.

6.

7.

8 .

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