NASA / TMn1999-209089 AIAA-99-2730

An Air-Breathing Launch Vehicle Concept for Single-Stage-to-Orbit

Charles J. Trefny Glenn Research Center, Cleveland, Ohio

Prepared for the 35th Joint Propulsion Conference and Exhibit sponsored by the AIAA, ASME, SAE, ASEE Los Angeles, California, June 20-23, 1999

National Aeronautics and Space Administration

Glenn Research Center

May 1999 Acknowledgments

The Trailblazer team at the NASA Glenn Research Center performed the work presented herein. This team includes personnel from the Turbomachinery and Propulsion Systems Division, the Engineering Design and Analysis Division, the Propulsion Systems Analysis Office, the Hypersonics Projects Office, and the Launch Vehicle and the Transportation Projects Office. The Hypersonic Airbreathing Propulsion Branch, and the Systems Analysis Office at the NASA Langley Research Center, and the Advanced Space Transportation Program Office at the NASA Marshall Space Flight Center provided consultation and review. This work is supported by the NASA Propulsion Base Research and Technology Program.

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NASA Center for Aerospace Information National Technical Information Service 7121 Standard Drive 5285 Port Royal Road Hanover, MD 21076 Springfield, VA 22100 Price Code: A03 Price Code: A03 AN AIR-BREATHING LAUNCH VEHICLE CONCEPT FOR SINGLE-STAGE-TO.ORBIT

Charles J. Trefny* National Aeronautics and Space Administration Glenn Research Center Cleveland, Ohio 44135

Abstract propellant mass fraction of 78.4 percent and an overall oxidizer-to-fuel mass ratio of 2.80. Based on The "Trailblazer" is a 300-1b payload, single-stage- preliminary structural modeling, the gross lilt-off- to-orbit launch vehicle concept that uses air-breathing weight is 130,000 lb. Numerous technical issues are propulsion to reduce the required propellant fraction. being addressed in a propulsion technology maturation The integration of air-breathing propulsion is done program. considering performance, structural and volumetric efficiency, complexity, and design risk. The resulting Introduction configuration is intended to be viable using near-term materials and structures. The aeropropulsion perform- A long-term NASA goal is to reduce the cost of ance goal for the Trailblazer launch vehicle is an space access. One means of achieving this goal is to equivalent effective specific impulse (I*) of 500 sec. develop reusable launch vehicles. Single-stage-to-orbit Preliminary analysis shows that this requires flight in (SSTO) vehicles are one class in a larger group of the atmosphere to about Mach 10, and that the gross concepts under consideration. NASA's X-33 program lift-off weight is 130,000 lb. The Trailblazer will provide information on the near-term viability of configuration and proposed propulsion system SSTO rockets, and the use of advanced, lightweight operating modes are described. Preliminary perform- materials in launch vehicle structures. The present work ance results are presented, and key technical issues are will advance the understanding of the use of air- highlighted. An overview of the proposed program plan breathing propulsion. Air-breathing propulsion reduces is given. the propellant fraction required for SSTO. However, the overall benefit in terms of SSTO feasibility, design Summary margin, and cost reduction will depend on weight, and complexity. Engineering design, fabrication, and The Trailblazer is a reusable, single-stage-to-orbit demonstration is required to accurately evaluate these launch vehicle concept, intended to reduce the cost of factors. Therefore, emphasis is placed on the space access by making optimum use of air-breathing engineering of a particular concept called "Trailblazer" propulsion. The Trailblazer development program is (Fig. !), that has good potential for structural based on maturation of a 300-1b payload reference efficiency, and minimum complexity. Trailblazer is vehicle to a point at which commercial development intended to make optimum use of air-breathing could proceed. The propulsion system operates in four propulsion tot the acceleration to orbit mission. modes including ramjet, scramjet, and rocket modes from lilt-off to orbit. A new low speed mode, intended The proposed program (Fig. 2) is based on to minimize weight and complexity is proposed. maturation of the Trailblazer "reference vehicle" design to a point at which commercial development could Preliminary analysis of the conceptual design indicates begin. The design will evolve based on results from a that the inlet capture area is sufficient for air-breathing propulsion technology maturation program, and a flight acceleration to Mach 10, resulting in a required demonstration. The propulsion technology maturation consists of complementary experimental, numerical, *Senior Member-AIAA.

NASAFFM--1999-209089 1 andanalyticalefforts.Plannedexperimentsrangefrom nacelle and boundary-layer diverter surfaces are proof-ofconceptandcomponentdevelopmenttests,to accounted for in the net thrust. flight-likepropulsionsystemtests.CFDanalysisisused to guidecomponentand testrig design,estimate AV Change in velocity performance,andanalyzeexperimentalresults.The purposeof flight demonstrationis to validate W Instantaneous vehicle weight performanceandstructuralweightestimates,anddrive allmanufacturingandoperationalissuestoresolution. W Inert mass fraction, vehicle weight at lift-off minus payload and ascent propellant Thepurposeof this paperis to introducethe Trailblazerreferencevehicleconceptualdesign,andthe et Angle of attack rationalebehindconfigurationchoices.Thepropulsion configurationand thermodynamiccycleswill be T Angle between flight path and Earth's surface. described.Preliminarysystemperformanceresultswill bcpresentedandtechnicalissueshighlighted. System Performance Goals

__mbols The equivalent, effective specific impulse (I*) is used herein as a measure of "aeropropulsion" perform- D Total drag fl)rce includingall surfacesnot ance. A brief explanation of I* follows; a complete accountedforin thenetthrust.Actsin directionof derivation can be found in Ref. 1. The effective specific flight. impulse (I,) is defined as the sum of all forces in the direction of flight due to propulsion, aerodynamics and g_ Proportionalityconstant,32.174lb,,,-ft/lbFsec"_ gravity, divided by the propellant flow rate. I, for an air-breathing launch vehicle can vary by an order of H Altitude magnitude during ascent, rendering its use in the traditional rocket equation invalid. I* is the constant, fell Effective specific impulse. The sum of all forces "equivalent" value of I, that, when used in the rocket acting on a flight vehicle in the direction of flight, equation, results in the correct mass ratio. I* therefore divided by the propellant flow rate represents the efficiency at which a launch vehicle expends its propellant to achieve a given velocity. I..... Specific impulse based on net thrust Figure 3 shows the effect of I* on the propellant I._. Specific impulse based on vacuum thrust in rocket fraction required (PFR) to achieve SSTO. At I* values mode (mode IV) approaching the maximum theoretically possible for chemical rockets, the PFR to achieve SSTO is about I* Equivalent, effective specific impulse. The 90 percent of the gross lift-off weight (GLOW). The constant value of specific impulse that, when used development of an SSTO rocket thus depends on in the rocket equation, results in the correct mass reducing the inert mass fraction to about 10 percent. ratio for a launch vehicle with variable effective Launch cost reduction depends on reusability of which spccific impulse SSTO is only one component. Another is increased design margin, which is in conflict with the L Aertxlynamic lilt force acting in direction normal requirement for minimum inert mass. The prospect of to flight path an air-breathing launch vehicle provides a means of increasing I* well beyond that of chemical rockets. The M,, Free-stream Mach number associated reduction in propellant fraction does not, however, guarantee SSTO closure at a practical P Rocket chamber pressure GLOW, increased margin for reusability, or reduced operations and maintenance costs. Mitigating factors q,. Free-stream dynamic pressure must be considered. These include the weight and complexity of air-breathing components, the burden of T Net axial thrust force. Based on a control volume flight within the atmospherc (heating, drag, and defined by the captured airstream and nacelle structural loading), and a lower propellant bulk density surfaces, extending from the inlet spike leading (due to the exclusive use of hydrogen in air-breathing edges to thc vehicle trailing edge. Forward-facing

NASA/TM--1999-209089 2 modes)that tends to increasetank weightan shapes. Vertical li_off also eliminates many safety aerodynamicdrag.A systemsapproachtoair-breathing issues associated with high-speed taxi. The propellants launch vehicledesign is requireddue to the are liquid oxygen (LOX) and liquid hydrogen (LH,). interrelationshipof thesefactorswithaeropropulsionThe energy per unit mass and heat capacity of hydrogen performance. wcrc deemed necessary. Finally, the reference vehicle is designed to carry 300 Ib to low-Earth orbit. This is to Thenetbenefitof higherI* dependsonthedegree minimize the scale and cost of a relevant demonstration to whichthemitigatingfactorsdepletethestructural vehicle. The design requirements allow subsequent marginaffordedbyincreasedaeropropulsionperform- scaling to larger payloads without regard to runway ance.Thisisdepictedin Fig.4 wheretheinertmass length or weight restrictions. fractionis plottedas a functionof I* tbr various payloadfractions.A netsystembenefitcanresultonly Vehicle Configuration and Propulsion Integration if the vectorsumshowntendstowardincreasing The basic vehicle configuration is pictured in margin.The Trailblazerprogramwill determine Fig. 5. The dimensions shown are preliminary and whetheror not the applicationot" near-termair- correspond to a GLOW of 130,000 lb. Three semi- breathingpropulsiontechnologyto an SSTOlaunch circular propulsion pods are mounted symmetrically at vehiclewill resultin sufficientmarginandreusability 120 ° intervals around the periphery of the vehicle's to warrantcommercialdevelopment.The goalfor circular cross section. This arrangement provides for initialdemonstrationhasbeensetatI* equalto500sec, diversion of the forebody boundary layer and results in andan inertmassfractionof 20 percent(1 percent an axial total thrust vector. The forebody is parabolic payloadfraction).Thesevaluesare lessthanthat with a 10° half-angle nose. This shape was chosen for theoreticallypossibleonboththeaeropropulsionand high structural and volumetric efficiency and for low structuralweightaxesof Fig.4, butmayrepresenta drag without regard for inlet precompression. The inlet nearer-term,morepracticalsolutionto costreduction spikes are offset from the forebody surface on thantheSSTOrocket,andhigherpertormanceair- boundary-layer diverter pylons. As the conceptual breathingconceptssuchastheNationalAero-Space design evolved, the required tank volumes and Plane(NASP).-'Thistargetrepresentstheleastrisk,and propulsion system length placed the leading edge of the leavesroomforevolutionwithsubsequentadvancesin diverters tbrward of the tangency point on the parabolic propulsion,materials,andactively-cooledstructures. forebody. The aft-facing projected area of the vehicle is used for nozzle expansion, resulting in an exit-to- Trailblazer Reference Vehicle Conceptual Design capture area ratio of 2.78:1. At low speed, this integration is intended to provide an altitude- The Trailblazer conceptual design philosophy is to compensating effect, similar to that of a plug nozzle, take optimum advantage of air-breathing propulsion that minimizes over-expansion losses. The wings are performance. That is, not to exceed the point of sized such that the total plan|orm generates sufficient diminishing returns on increasing I*. The basic con- lift to support the vehicle at a maximum angle-of-attack figuration and operational scheme has been constrained of 6° in air-breathing modes. to provide high potential for lightweight, durable structures and tanks, and to minimize risk and The aeropropulsion performance of the vehicle is complexity. Axisymmetric geometry provides inherent strongly dependent on the inlet capture area. The structural efficiency, and lower uncertainty in maximum viable air-breathing Mach number and IL.,,are aerodynamic and structural design, leading to a functions of air capture. The optimum capture area reduction in required factors of safety. cannot be determined, however, until the structural weight and thermal protection characteristics of both Design Requirements the vehicle and propulsion system are better The basic configuration and operational schemes understood. The current total capture area of 98.41 ft -_ were chosen to provide good contrast to prior air- (69.3 percent of the forebody maximum cross-sectional breathing launch vehicle concepts such as the NASP. It area) is sized for a maximum air-breathing Mach is required that the vehicle be SSTO and reusable. The number of about l(J. number of missions is not yet defined. Vertical lift-off and horizontal landing are proposed to minimize wing Although the forebody is not intended to be a structure and landing gear weight. This also minimizes compression surface, flow-field calculations _ have reliance on aerodynamics at low speed, which can drivc revealed that significant compression does persist to the horizontal take-off systems to non-optimum structural diverter leading edge station resulting in significantly

NASA/TM--1999-209089 3 greatermasscapturethanthatofafreestreaminlet.At numbers. It is also the point at which the centerbody Mach10,thestaticpressure(stream-thrustaveraged contacts the cowl contour. The current geometry results overthecrosssectionof thecapturedstream)istwice in a contraction ratio of 15 with shock-on-lip at that in the freestream,at a contractionratio of Mach 6. Inlet contours and performance estimates about1.5. based on axisymmetric, turbulent, Navier-Stokes calculations are presented in Ref. 3. The throat is Theinletsmustoperatewhenthevehicleis atan angled toward the inlet axis at 15° to reduce the length, angle-of-attack. Three-dimensional,turbulent, weight and wetted area of this portion of the flow path. ParabolizedNavier-Stokescalculationshavebeen The 12° spike angle is also meant to minimize length. performedon theforebodyshapeat Machnumbers The maximum spike and throat angles for acceptable from1.2to10,andangles-of-attackof0°,3°,6°,and9°. inlet aerodynamic operability without boundary layer Thesesolutionsindicatethattheflowfieldatthespike bleed is one of the key inlet design issues. The leadingedgeplaneisfreeof severedistortionswhen incorporation of a boundary-layer bleed system is to be theangle-of-attackislessthanorequalto6°. Isolated, avoided due to the weight and complexity of the axisymmetricmixed-compressioninlets havebeen associated ducting, metering, and control systems. operatedatangles-of-attackgreaterthan7°.4Additional numericalandexperimentalmodelingwill berequired Rocket Element. There is one rocket element per to determinetheexactoperatinglimitsof thepresent flow path, located in a semi-circular hub that is fixed configuration,whichis in closeproximityto the with respect to the vehicle. The centerbody translates forebody. over the hub, and their trailing edges are coincident when the centerbody is in its aft-most position. The Thevehiclecrosssection(Fig.6)revealsthesize total duct cross-sectional area at the hub trailing edge andlocationofthepropellanttanks.Thesearesizedto station is 0.4 times the inlet capture area. The rocket accommodatetherelativevolumesof LOXandLH2 element operates at a constant oxidizer-to-fuel mass required. The LOX tank was placed forward of the LH 2 ratio (O/F) and variable chamber pressure. tank to increase longitudinal static stability. However, this is not the preferred arrangement for minimum The O/F ratio must be optimized during reference structural weight. Other more optimum arrangements vehicle maturation, considering rocket It, and may be possible once stability requirements are better propellant density, which affects vehicle weight and understood. drag. A baseline value of 6 has been used in the present analysis. A maximum chamber pressure of 1500 psia Propulsion System Description was chosen as a baseline value to be optimized during Figure 7 is a cut-away view of a propulsion "pod." definition of the propellant cycle and cooling circuit Its axisymmetric design provides the potential for good design. The throat area is 0.0068 times the inlet capture structural efficiency and minimizes aerodynamic design area which results in a lift-off thrust-to-weight ratio of and analysis uncertainty due to the two-dimensional about 1.7 for a 130,000 lb GLOW. The rocket element nature of the flow field. The axis of symmetry is a line expansion ratio is 10, constrained by the available hub parallel to the vehicle axis at a distance equal to the cross section, which is 8.5 percent of the inlet capture diverter radius from vehicle axis. area. The hub cross section must not be excessive because it acts as a base area within the flow path when Inlet. The inlet is of translating-centerbody, mixed- the rocket is not operating. compression design. Centerbody translation allows starting, provides variable contraction ratio, and closes The rocket element's plug nozzle configuration off the duct for rocket-mode operation. This provides pressure compensation as the rocket is configuration has minimum sealing, support, and throttled and ram pressure increases during low speed actuation issues. At full retraction, the spike and propulsion mode. Its semi-circular cross-section also diverter planforms are coincident. When extended, the integrates well with the semi-circular hub base. It is edges of the spike overhang the diverter pylon. The designed with 50 percent internal expansion to allow inlet duct cross-section is semi-annular. It is not strictly throttling to approximately 20 percent of maximum axisymmetric though, since the cylindrical surface upon thrust. There are numerous issues with throttling and which the annulus terminates is not a radial plane of cooling of annular rockets that are currently being symmetry. A step in the centerbody contour at the addressed. The pertbrmance increments gained by throat is intended to provide a degree of inlet isolation, rocket throttling may be outweighed by increases in and a surface for axial fuel injection at high flight Math weight and complexity.

NASA/TM--i 999-209089 4 Ramjet Duct and Nozzle. Aft of the rocket element, increases complexity and may have an adverse effect a conical cowl surface extends downstream at a 5.67 ° on other modes of operation. wall angle until intersected by a 15° half-angle "expansion cone." The surface of this virtual expansion At lift-off, the rocket operates at maximum cone defines the trailing edge geometry of the cowl and chamber pressure. The inlet centerbody is fully vehicle surfaces. The expansion cone apex is located extended. The open inlet ventilates the ramjet duct such that the maximum internal area of the flow path is preventing over-expansion of the rocket. As the vehicle 1.25 times the inlet capture area, and the distance gains speed, the airstream can be fueled and burned to between the hub and this area is 2.85 times the cowl lip generate ramjet thrust. As the proportion of ramjet radius. The maximum cross-sectional area is sized to thrust becomes significant, the rocket chamber pressure accommodate stoichiometric combustion of the can be reduced and optimized for maximum l f,. The expected airstream at low supersonic flight Mach high thrust rocket cycle gives way to the high specific numbers. The duct length must be sufficient for impulse ramjet cycle in this manner. When ramjet completion of the ramjet combustion process. These are thrust alone is sufficient, the rocket is shut off baseline values and will be revised as higher fidelity completely, defining the transition to mode II. analysis and test data become available. This crude method of generating the nozzle contour will also be A two-stream, one-dimensional model of this cycle revisited. It does, however, illustrate desirable has been developed + to determine the optimum ramjet integration and altitude-compensation features. The fuel-air ratio and thermal throat location lbr a given total area ratio from the rocket element throat to the flight Mach number and rocket chamber pressure. vehicle aft projected area is 409:1. Pertbrmance maps generated using this procedure then allow the rocket chamber pressure to be varied during Propellant System. Conceptual design of a power- trajectory optimization for maximum I*. balanced propellant delivery system including turbopumps, valves, heat exchangers (actively cooled Combustion of the airstream and the resulting surfaces) and associated piping and controls is in thermal throat must occur at a cross-sectional area on progress. Various propellant cycles will be examined to the order of the inlet capture area to avoid sub-critical determine those best suited to the present application. operation of the inlet and associated spillage drag. One Some factors to be considered are the large areas of method to accomplish this is the use of in-stream spray actively cooled surface that will lead to high pressure bars such as those reported in Refs. 7 to 9. Spray bars drops, and may require more fuel for cooling than that are impractical for the present problem however, required for combustion in air-breathing modes. Also, because they would have to be retracted during independent control of fuel and oxidizer flow rates is operation in modes III and IV to avoid severe required in the various propulsion modes described expansion process losses and cooling problems. The below. present approach is to use fuel injectors in the annular inlet diffuser to distribute fuel in the airstream. The Propulsion System Operating Modes premixed stream is then ignited by its confluence with the rocket, and the combustion process proceeds The Trailblazer propulsion system operates in four downstream to completion at a large cross-sectional distinct modes during the earth-to-orbit ascent. The area. Fuel is distributed in the annulus such that thin, flow path is designed to accommodate the noncombustible layers exist at the walls to prevent thermodynamic process of each. flashback. Some mixing between the rocket stream and this "buffer layer" must then occur for ignition to Mode I--Lift-off and Low Speed proceed. This mode, depicted in Fig. 8(a) provides the high thrust required for lift-off, and initiates the high specific Peak performance is obtained with the thermal impulse ramjet cycle. The rocket and air streams throat at a specified cross-sectional area that varies with interact along a matched-pressure boundary but are not flight Mach number. The axial location and theretore intended to mix. This mode of operation has been cross-sectional area of the thermal throat depends on dubbed the independent ramjet stream ORS) cycle. the point of ignition and the rate at which the flame Traditional ejector-ramjct cycles _ have slightly higher travels across the duct. The radial distribution of fuel thermodynamic perlbrmance, but would require mixing can therefore be used to control thermal throat location enhancement devices or multiple rocket chambers to based on pressure sensed at specific locations in the shorten the length of duct required for mixing. This inlet.

NASA/TM--1999-209089 5 Therearea numberof issuesassociatedwiththis The issues associated with mode III are those schemethat are currently under investigation. The two- generally associated with scramjet propulsion (see dimensional features of flame propagation into a Ref. 10). In addition, the nozzle contours are non-ideal thermally-choking stream are unknown. Auxiliary due to the multimode nature of the flow path. The flame-holding sites may be required to avoid an effective specific impulse is very sensitive to the excessively long duct. Whether or not sufficient control expansion process efficiency, especially approaching of the thermal throat location is possible must be the expected maximum air-breathing Mach number of determined. The prevention of premature combustion about 10. due to flashback must be assured, especially in the presence of realistic inlet flow distortions. Should the Mode IV--Rocket present scheme prove infeasible, a retractable spray bar Scramjet thrust per unit airflow decreases with arrangement could be considered. Another option is to flight Mach number. As I+,, approaches that of the revert to the simultaneous mixing and combustion rocket vacuum specific impulse (I,,) or the flight Mach cycle, described in Ref. 5, where the airstream is fueled number approaches a constraint based on system by mixing with fuel-rich rocket effluent. In this cycle, considerations, transition to mode IV occurs. The the thermal throat location could be modulated by transition sequence begins as the vehicle pitches up, active control of mixing intensity. Multiple rocket and climbs to a minimum dynamic pressure of about chambers, or mixing enhancement devices, would 500 psfa at constant Mach number. The centerbody is likely be required to avoid excessive duct length. then translated aft to close off the air-breathing flow path and the rocket is reignited (Fig. 8(c)). Mode II--ThermaI-Throat Ramjet When sufficient ramjet thrust is available, the The potential for high performance exists due to rocket is shut off. The started inlet requires a specified the large area ratio from the rocket throat to the vehicle exit pressure Ibr optimum performance. As in mode I, aft-projected area. The area distribution is not ideal, this requires control of the thermal throat location. If however, with a significant gap between the rocket exit the transition Mach number is below about 2.5, a fuel area and the air-breathing duct. This "free-expansion" distribution and combustion process similar to that used must be managed by pressurizing the cavity upstream in mode I would be required, except that piloting would of the rocket with a small amount of bleed flow. The be from the recirculation zone formed downstream of efficiency of this type of process has been studied the rocket hub. Some form of auxiliary piloting may parametrically for axisymmetric geometries in Ref. I 1. also be required. As the free stream stagnation The sensitivity of I* to mode IV pertbrmance is high, temperature increases, auto-ignition and flashback to roughly 0.7 sec of 1" per sec of I+_, since about the point of injection would be unavoidable. However, 60 percent of the total velocity is imparted to the the required thermal throat cross-section is also reduced vehicle in this mode. Uncertainty in 1 must be so that the mixing-limited combustion process pictured minimal. Three-dimensional modeling of the expansion in Fig. 8(b) is feasible. Fuel injection from the walls at process is required. The rocket must bc throttled in various axial stations is used to control the thermal mode IV if a maximum acceleration constraint is throat location. The Mach number at which the mixing- imposed. limited ramjet combustion process is feasible has not yet been determined. Techniques for piloting the To protect the cowl lips during reentry, inert gas premixed combustion process and controlling the would be bled into the forward cavity as pictured in thermal throat location must be examined. Fig. 8(d).

Mode Ill--Supersonic Combustion Ramjet A-nal_ Between approximately Mach 5 and 6, the combination of inlet losses, hig.h duct pressure, cooling Trajectory Optimization requirements, and non-equilibrium chemistry make The aeropropulsion performance of the reference transition to the supersonic combustion mode vehicle is determined by trajectory optimization using beneficial. Fuel is injected axially from the centerbody the Optimal Trajectories by Implicit Simulation step, and at various downstream stations, to tailor the (OTIS) '_ program. Inputs to OTIS include propulsion combustion distribution for optimum performance and performance, configuration aerodynamics, various avoid choking. The inlet contraction ratio can be constraints, and the required orbit. OTIS determines the optimized as information on weight and cooling trajectory that tnaximizes the final weight and therefore becomes available. I*. This d(ees not necessarily represent the system

NASA/TM--1999-209089 6 optimumthough.Trajectoryoptimizationispartof an ramjet thrust becomes available. Full transition to iterativeprocessthataccountsfor theeffectsof the ramjet mode occurs at Mach 2.04. The minimum Mach variousconstraintsonstructuralweight. number at which this transition can occur depends on air capture (both inlet area, and inlet mass flow ratio). Mode1 propulsionperformancewascalculated The effect of increasing the constrained minimum usingthemethodof Ref.6.Netthrustandpropellant chamber pressure is being evaluated in light of the flowratesweredeterminedasfunctionsof freestream additional complexity introduced by deep throttling. To Machnumberand rocketchamberpressure.The satisfy the 4g maximum acceleration constraint, it is optimumscheduleof rocketchamberpressureis necessary to throttle the rocket to 625 psia in mode IV. determinedconcurrentlywith theoptimumtrajectory. The vehicle angle-of attack and lift-to-drag ratio along Performancein modesII andIII wascalculatedusing the optimum trajectory appear in Fig. 1 !. Just alter lift- theRamjetPerformanceAnalysis(RJPA)program." ofl\ a pitch-over maneuver initiates transition to the Inletrecoveryanddragsfor all air-breathingmodes more horizontal flight path required for air-breathing werebasedon the 2-D axisymmetriccalculations flight. The angle-of-attack peaks at just over 4 ° after an describedin Ref.3.ThemodeIV I wasestimatedto initial 25° negative spike. The angle-of attack remains be 450 sec usingthe methodof Ref. 11. An below 4° from pitch-over to Mach 10, when the 6° atmosphericback-pressuretermissubtractedfromthe constraint is met during the climb prior to transition to resultantthrust.Theaerodynamiccharacteristicsof the mode IV. The vehicle thrust-to-weight ratio, and vehicleasfunctionsof Machnumberandangle-of- flightpath angle on the optimum trajectory are shown in attack were evaluatedusing the Aerodynamic Fig. 12. The pitch-over, and transition to nearly PreliminaryAnalysisSystem'_(APAS). horizontal flight are evident in the flightpath angle trace. The thrust-to-weight ratio trace indicates that the Theconstraintslistedin TableI wereimposed acceleration remains well below the 4g limit throughout duringtrajectoryoptimization.Themaximumdynamic modes I to III. However, rocket throttling is required in pressureandmodeII1 to IV transitionMachnumber mode 4 to remain within the acceleration constraint. constraintswill be revisitedoncethe sensitivityof The net propulsive specific impulse is shown in Fig. 14, vehicleweightto theseparametersis understood.An along with the effective specific impulse. The easterly,verticallaunchfrom 28.5° latitude,to a difference between the two curves represents the 220nmicircularorbitwasassumed. retarding forces of vehicle drag and gravity. Transition from mode I to II is marked by a sharp increase in I, as A summaryof key points along the optimum the rocket is shut off. Transition from mode III to IV trajectory is presented in Table II. Time spent in occurs at Mach 10 as constrained, although the net atmospheric flight in air-breathing propulsion modes is thrust is sufficient for continued acceleration in less than 6 min. Therefore, thermal protection, and mode III (IL.,,is greater than that of mc_e IV). This cooling system designs must be based on transient potential for higher 1" must be carefully traded against heating analysis. Total elapsed time from lift-off to reduced propellant bulk density and increased heat orbit is 39 min, 48 sec. The final mass in orbit is load. 21.6 percent of the GLOW (78.4 percent PFR). I* for this trajectory is 509.3 sec, based on the initial and final Preliminary Closure Results vehicle weight, and the total change in inertial velocity The trajectory shown in Fig. 9 defines the thermal, due to propulsion. The overall vehicle O/F for this and mechanical loads, the propellant fraction required, trajectory is 2.80. and relative size of the propellant tanks. Structural design and optimization based on these inputs is in The optimum trajectory appears in Fig. 9. It is progress. Scaling of a preliminary structural weight characterized by flight at maximum dynamic pressure model indicates that the vehicle propellant fraction to the maximum air-breathing Math number of 10. At equals that required (the vehicle is "closed") at a gross this point, the vehicle climbs to the minimum dynamic lilt-off weight of about 130,000 lb. The inert mass pressure of 500 psfa, transitions to mode IV, then fraction is slightly greater than the target value of follows a higher altitude path more optimum for rocket 20 percent resulting in a payload fraction of propulsion. The remaining mode IV powered phase, 0.23 percent. coast phase, and circularization burn are not shown in the figure. The optimum chamber pressure (Fig. 10) is Preliminary structural weight modeling is based on the maximum from lift-off through the transonic drag nonintegral, graphite-epoxy composite propellant tanks rise, followed by a sharp reduction as higher efficiency that account for about 7 percent of the dry weight.

NASA/TM--1999-209089 7 Actively-cooledcarbon-carbon,and carbon-silicon- Ultimately, a muitidisciplinary optimization, such carbidecompositeswere assumedfor propulsion as that presented in Ref. 15 could be perJbrmed. The systemsurfaceswith titaniumandaluminum-lithiuminter-relationship between aeropropulsion performance, substructures.Constructiontechniquesand com- structural weight and complexity dictates that the entire patibilityissuesarebeingaddressed.Thepropulsion launch vehicle system is considered during conceptual systemweightis about39percentof thedryweight. design. The vehicleaeroshelland wing structuresconsist largely of carbon-carbonsandwichpanelswith A series of propulsion component test rigs are haifnium-carbidecoatingandmakeupabout30percent currently under development. These will address the of thedryweight.Allowancesweremadeforregions operability issues discussed herein, and provide requiringactive-coolingor additionalpassivethermal performance validation. Numerous numerical and protection.The useof halfnium-diborideis being analytical studies are also in progress to optimize, and studiedtbrtipandleadingedgeareas. increase the fidelity of the reference vehicle design.

Thecenter-of-gravityhasbeencalculatedas a Acknowledgments functionof theamountsof LOX and LH, remaining, using a mass distribution based on the preliminary The Trailblazer team at the NASA Glenn Research model. The vehicle is statically stable in the pitch axis Center performed the work presented herein. This team during ascent, but requires some tbrm of pitch control includes personnel from the Turbomachinery and for trim. Combinations of elevon deflection, differential Propulsion Systems Division, the Engineering Design throttling, and angle-of-incidence are currently being and Analysis Division, the Propulsion Systems studied. With empty propellant tanks, the center of Analysis Office, the Hypersonics Projects Office, and gravity moves aft and the configuration exhibits neutral the Launch Vehicle and the Transportation Projects static stability. Active pitch control may be required tor Office. The Hypersonic Airbreathing Propulsion landing. Branch, and the Systems Analysis Office at the NASA Langley Research Center, and the Advanced Space Concluding Remarks Transportation Program Office at the NASA Marshall Space Flight Center provided consultation and review. A configuration with potential for structural This work is supported by the NASA Propulsion Base efficiency and simplicity can meet the 500 sec I* Research and Technology Program. aeropropulsion performance goal with a maximum air- breathing Mach number of about 10. The inlet capture References area required for acceleration to Mach 10 is not prohibitive even though the forebody is not optimized I. Escher, W.J.D.; Hyde, E.H.; and Anderson, D.M.: for air-capture. The vehicle O/F is 2.80 under the cycle A User's Primer for Comparative Assessments of performance assumptions stated, which include the IRS All-Rocket and Rocket-Based Combined-Cycle low-speed cycle, and a rcx:ket I of 450 sec. The Systems for Advanced Earth-to-Orbit Space vehicle aerodynamics provide sufficient lift at less than Transportation Applications. AIAA Paper 6° angle-of-attack during air-breathing modes. Based 95-2474, July 1995. on preliminary structural architecture and weight 2. Barthelemy, R.R.: The National Aero-Space Plane modeling, the GLOW is 130,000 lb. Program. AIAA Paper 89-5053, July 1989. 3. DeBonis, J.R.; Trefny, C.J.; and Steffen, C.J.: These preliminary results warrant continued Inlet Development for a Rocket Based Combined maturation of the reference vehicle concept. Current Cycle, Single Stage to Orbit Vehicle Using plans call for more detailed design of the propulsion Computational Fluid Dynamics. AIAA Paper system including actively cooled composite structures 99-2239, June 1999. and propellant delivery systems, to refine weight 4. Choby, D.A.: Tolerance of Mach 2.50 estimates, and assess complexity and cost savings Axisymmetric Mixed-Compression Inlets to potential. System trades on a number of parameters, Upstream Flow Variations. NASA TM X-2433, such as rocket O/F ratio and chamber pressure, January 1972. maximum air-breathing Math number, maximum dynamic pressure, inlet capture area, and material selection await more detailed weight modeling.

NASA/TM--1999-209089 8 5. Biilig,F.S.:TheIntegrationof theRocketwith 11.Smith,T.D.; Steffen,C.J.; Yungster, S.; and theRam-Scramjetasa ViableTransatmospheric Keller, D.J.: Performance of an Axisymmetric Accelerator.I1" InternationalSymposiumon Rocket Based Combined Cycle Engine During Air-Breathing Engines, Paper 93-7017, Rocket Only Operation Using Linear Regression September1993. Analysis. NASA TM-1998-206632, March 1998. 6. Yungster,S.;Trefny,C.J.:Analysisof a New 12. Hargraves, C.R.; and Paris, S.W.: Direct Rocket-BasedCombined-CycleEngineConcept Trajectory Optimization Using Nonlinear atLowSpeed.AIAAPaper99-2393,June1999. Programming and Collocation. Journal of 7. Kerslake,W.R.: Combustionof Gaseous Guidance, Vol. 10, No. 4, July-August 1987. HydrogenatLowPressuresin a 35° Sectorof a 13. Pandolfini, P.P.; and Friedman, M.A.: Instructions 28-inch-DiameterRamjetCombustor.NACARM for Using Ramjet Performance Analysis (RJPA) E58A21a,April1958. IBM-PC Version !.24. The Johns Hopkins 8. Krull,G.H.;andBurley,R.R.:Effectof Burner University Applied Physics Laboratory, JHU/APL DesignVariableson Performanceof 16-inch- AL-92-PI75, June 1992. DiameterRam-JetCombustorUsingGaseous- 14. Cruz, C.I.; and Willhite, A.W.: Prediction of Hydrogen Fuel. NACA RM E56J08, High-Speed Aerodynamic Characteristics Using January1957. the Aerodynamic Preliminary Analysis System 9. Cervenka,A.J.;andSheldon,J.W.:Methodfor (APAS). AIAA Paper 89-2173, July 1989. ShorteningRam-JetEnginesby Burning 15. Olds, John R.: Results of a Rocket-Based HydrogenFuelin theSubsonicDiffuser.NACA Combined-Cycle SSTO Design Using Parametric RME56G27,October1956. MDO Methods. SAE Paper 941165, April 1994. 10.Heiser,WH.;andPratt,DT.:HypersonicAirbreathing Propulsion.AmericanInstituteof Aeronauticsand Astronautics,WashingtonD.C.1994.

TABLE I.--TRAJECTORY OPTIMIZATION CONSTRAINTS Parameter Enforced Minimum Maximum value value

Free stream dynamic pressurc (psfa) All modes 500 1500 Angle-of-attack (deg) Modes 11, III -6 +6 Rocket chamber pressure (psia) Modes I, IV 300 1500 Total acceleration (g's) All modes 4 Mode Ill-IV transition Mach number Mode Ill 10

TABLE II.--TRAJECTORY SUMMARY Event Elapsed Altitude Ineoial Flight path Free-stream Vehiclc time, velocity, angle, Mach weight, min:sec f_sec deg nmnber Ibm Lift-off 1.340 9O 130,000 Mode 1-2 transition 1:07 34,399 fl 3,321 12.9" 2.04 94,630 Mode 3-4 transition 5:45 126,704 fl 11,510 3.1 9.85 81,915 28.926 Begin coast 8:31 45.02 nrni 26,018 1.6 .... 28,926 Begin circ. burn 39:42 221.60 nhfi 24,746 0 .... End circ. bum 39:48 221.62 nmi 25,146 0 .... 28,138

NASA/TM--1999-209089 9 0.3

air.breathing "Trailblazer = reference

vehicle goal -_.....

0.2

" Structural weight 3; 0.1 _ _ burden

c_ °_ Aero-propulaion

_ o_ _ I_rfoml_nce

_o_¢ o benefit

0.0 300 400 500 600 700

Equivalent Effective Specific Impulse, t" (Ib_-sec.,/tb_)

Figure 4.--SSTO launch vehicle design space.

i T

Figure 1,--Trailblazer reference vehicle.

114.688 Reference Veh0cle Concept Commercial Init6al Feas=bility S_udy Matural_on Development

lo_ a teusa_e. SSTO RGCC- Re<_Cl uflcet_intles in slrUCluraJ Inlt_ v_le based onret_te_ce proplllkld launci_vlll_cle mass Iract_c,n_d conCepl conCepl Begin _¢_u@on ot I.....the "Refece¢lcl"...... Veh_e" te_abihl_ Io a_ _CC_l_l_ level Hpedormarce and cOSlred_J "\ / \ Propulsion Technology Propulsion Syslem Fkght Maturation Oernonstralton

_ev elop _ vak_ propu_slo_ Val_d_tesystem pedorm_nce _d and _eh_c_econcedls thr_gh operability _naJlmodes _llh a sub- _o_n_ lest and a,'_aly_sot onblal sub-scale "Tmi_4_zee _oput_o_ d=mo _eh_e

I Figure 2._Pmposed program. i ..... 46.843 ---_

R7,470

diverter m radius _ _ /- R6.723 1.0

c,. 0.9 y radius

_- 0.8 5 76.256

0.7 ,Y

_. o.6 o SSTO RBCC Q. Syetetrm I 0.5 : i (AV = 25027 fps)

200 300 400 500 600 700 800 900 1000

Equivalent Effective Specific Impulse, I" (IbcsecJIb_)

Figure 3.--Effect of I* on propellant mass fraction for Dimensions in feet SSTO. Figure 5.--Reference vehicle geometry.

NASA/TM--1999-209089 10 Fuel injection and Ignition Flame Thermal premixing (Mo > 1) point front throat

/ / (a) Rocket exhaust

Started Fuel injection Thermal throat mixing / F LOx tank inlet and (ramjet mode) /: .___ // /-- LH 2 tank / ¢ /

\ 'l _ / Pressurization tank \ / \ / (b} _- pressurization tank L- Payload bay

Section A-A Inlet Cavity High _rea closed bleed ratio expansion Figure &--Preliminary vehicle layout. /: i

(c)

Inert gas bleed X. X

Hydrogen fuel F Trailing edge of injection sites _t_ / fixed hub containing \ / rocket elemen_

Translating Cowl I \ // \ / /- Plug nozzle centedood¥ --_ lip _ I \ / (d) \ X I \ / / . J-_ ', Figure 8.mPropulsion system operating modes. (a) Mode I, lift-off and low speed. (b) Modes II and ![I, ram and scramjet. (c) Mode IV, rocket. / i ' % L Diverter pylon / (d) Re-entry concept. L_ Ramjet duct and nozzle

Figure 7._Cut-away view of propulsion system.

NASA/TM--1999-209089 I I 160 45 90 Mode II(.IV ......

120 F_'slr_4lm dyni_l¢ ...**'" _" ...... _, 100 ...... _ ...... :::::---_:::::... ! •-/ Ttmu_ _ _" 80 "'"'" ...... "_"...... "...... :__:::::...... )_o _, " _' _o_= _ ce_l _eh so

40

_" f::::::::: ...... 05 10 oo 7 ...... o o 0 1 2 3 4 .5 6 7 8 9 10 11 12 0 1 3 4 5 6 7 S 9 10 11 12 Free-stream Mach number, M0 Free-stream Mach number, Mo Figure 12._Vehicle thrust-to-weight ratio and flight Figure 9.--Optimum trajectory. path angle on optimum trajectory.

1750 4O00

3500 -_ 1500 Q'_ 1250 _=300o _L' a_ 2500 1000 c_ ._=

_. 7so -5 E - 15017

"'" "-_) 1ooo

_') 50o rr

o 0 05 1 15 2 2.5 3 0 1 2 3 4 5 6 7 8 9 10 11 12

Free-stream Mach number, Mo Free-stream Mach number, Mo

Figure lO._Optimum chamber pressure schedule. Figure 13.--Specific impulse on optimum trajectory.

8 7

_tch_p _ io 6 51\ modi 1114V..

_ ',, 5E3

_ 2 -- Nlod_ ll_.IV

3_

"° 2

: ...... UI_. io4ra_ I 2 ,_

<-4 "_ "'_,,,.. "_ _

plltc h_ r I)_,gln .... o -8 -1 1 2 3 4 5 6 7 8 9 10 11 12

Free-stream Mach number, Mo

Figure 11 ._Angle-of-attack and li_to-drag ratio on optimum trajectory.

NASAJTM-- 1999-2[)9089 12 REPORT DOCUMENTATION PAGE FormApproved OMB No, 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED May 1999 Technical Memorandum 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

An Air-Breathing Launch Vehicle Concept for Single-Stage-to-Orbit

WU-523-61-23--00 6. AUTHOR(S)

Charles J. Trefny

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER National Aeronautics and Space Administration John H. Glenn Research Center at Lewis Field E-I1674 Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORTNUMBER

National Aeronautics and Space Administration NASA TM--1999-209089 Washington, DC 20546-0001 AIAA-99-2730

11. SUPPLEMENTARY NOTES

Prepared for the 35th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME. SAE, ASEE, Los Angeles, Calilornia, June 20-23, 1999. Responsible person, Charles J. Trefny, organization ccx:le 5880, (216) 433-2162.

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DIS_HIBUTION CODE

Unclassified - Unlimited Subject Category: Categories: 15 and 20 Distribution: Nonstandard

This publication is available from the NASA Center for AeroSpace Information. (301) 621-0390

13. ABSTRACT (Maximum 200 words)

The "Trailblazer" is a 300-1b payload, single-stage-to-orbit launch vehicle concept that uses air-breathing propulsion to reduce the required propellant fraction. The integration of air-breathing propulsion is done considering performance, structural and volumetric efficiency, complexity, and design risk. The resulting configuration is intended to be viable using near-term materials and structures. The aeropropulsion performance goal for the Trailblazer launch vehicle is an equivalent effective specific impulse (I*) of 500 sec. Preliminary analysis shows that this requires flight in the atmosphere to about Math I0, and that the gross lift-off weight is 130,000 lb. The Trailblazer configuration and proposed propulsion system operating modes are described. Preliminary performance results are presented, and key technical issues are highlighted. An overview of the proposed program plan is given.

14. SUBJECT TERMS 15. NUMBER OF PAGES 18

Reusable launch vehicles: Air breathing engines; Single phase to orbit vehicles 16. PRICE CODE A03

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified

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