AIAA 2002-5188 Technology Roadmap for Dual-Mode …mln/ltrs-pdfs/NASA-aiaa-2002-5188.pdf · AIAA 2002-5188 1 AIAA/AAAF ... fundamental research in supersonic combustion flow physics,

AIAA 2002-5188

1AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference

AIAA 2002-5188Technology Roadmap for Dual-Mode ScramjetPropulsion to Support Space-Access VisionVehicle Development

Charles E. Cockrell, Jr.Aaron H. AuslenderR. Wayne GuyCharles R. McClintonSharon S. WelchNASA Langley Research Center Hampton, Virginia, USA

11th AIAA/AAAF International Space Planes andHypersonic Systems and Technologies Conference

29 September – 4 October 2002 / Orleans, France

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Technology Roadmap for Dual-Mode Scramjet Propulsion to Support Space-Access Vision VehicleDevelopment

Charles E. Cockrell, Jr.*, Aaron H. Auslender†, R. Wayne Guy‡, Charles R. McClinton§, Sharon S. Welch§§

NASA Langley Research Center, Hampton, VA

Abstract

Third-generation reusable launch vehicle(RLV) systems are envisioned that utilizeairbreathing and combined-cycle propulsion to takeadvantage of potential performance benefits overconventional rocket propulsion and address goals ofreducing the cost and enhancing the safety ofsystems to reach earth orbit. The dual-modescramjet (DMSJ) forms the core of combined-cycleor combination-cycle propulsion systems for single-stage-to-orbit (SSTO) vehicles and provides most ofthe orbital ascent energy. These concepts are alsorelevant to two-stage-to-orbit (TSTO) systems withan airbreathing first or second stage. Foundationtechnology investments in scramjet propulsion aredriven by the goal to develop efficient Mach 3-15concepts with sufficient performance and operabilityto meet operational system goals. A brief historicalreview of NASA scramjet development is presentedalong with a summary of current technology effortsand a proposed roadmap. The technologyaddresses hydrogen-fueled combustor development,hypervelocity scramjets, multi-speed flowpathperformance and operability, propulsion-airframeintegration, and analysis and diagnostic tools.

IntroductionThe United States National Aeronautics and

Space Administration (NASA) has established astrategic goal of creating a safe, affordable highwaythrough the air and into space. Candidate third-generation reusable launch vehicle (RLV)architectures include single-stage and two-stageconcepts which utilize airbreathing, combined-cycleand combination-cycle propulsion systems to takeadvantage of potential performance gains overconventional rocket-propelled concepts. An access-

to-space roadmap has been established thatfocuses on airframe-integrated hypersonicairbreathing propulsion development throughfoundation technology investments, grounddemonstration and flight validation. Successfulimplementation of this roadmap requires a robusttechnology development program to mature aspectsof the propulsion system and integrated aero-propulsive vehicle performance through bothanalytic and experimental research.

Figure 1 shows a comparison of nominalspecific-impulse values for airbreathing enginecycles vs. rockets. The dual-mode scramjet (DMSJ)forms the core of combined-cycle or combination-cycle airbreathing propulsion systems and providesmost of the orbital ascent energy for single-stage-to-orbit (SSTO) airbreathing launch vehicle systems.The term “dual-mode scramjet” refers to an enginecycle that can operate in both subsonic combustionand supersonic combustion modes. Rocket-basedcombined cycle (RBCC) concepts are being studiedwhich integrate rocket thrusters with the DMSJflowpath for low-speed propulsion. Turbine-basedcombination cycle (TBCC) concepts are also beingexamined which integrate a gas turbine engine andDMSJ in a dual-flowpath configuration.

Hypersonic airbreathing propulsion researchconducted by NASA spans over 40 years.1-4

Historical work includes the hypersonic researchengine (HRE), airframe-integrated scramjet groundtesting and component development, the X-30National Aerospace Plane (NASP) program, and,more recently, the Hyper-X (X-43) flight

* Airbreathing Propulsion Program Manager, Advanced SpaceTransportation Program Office, Senior Member, AIAA.

† Assistant Head, Hypersonic Airbreathing Propulsion Branch.‡ Head, Hypersonic Airbreathing Propulsion Branch.§ Technology Manager, Hyper-X Program Office.§§ Head, Advanced Space Transportation Program Office.

Copyright © 2002 by American Institute of Aeronautics andAstronautics, Inc. No copyright is asserted in the United Statesunder Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimedherein for governmental purposes. All other rights are reservedby the copyright owner.

Turbojets

Scramjets

Ramjets

TurbojetsRamjet

Scramje

Hydrocarbon Fuels

Hydrogen Fuel

RocketRBCCTBCC

Mach Number10 20

Isp

Figure 1. Airbreathing Propulsion Performance.

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demonstration project. This work has also includedfundamental research in supersonic combustion flowphysics, analysis tools and diagnosticmethodologies.

Multidisciplinary research is required todevelop efficient airbreathing and combined-cyclepropulsion systems for airframe-integrated vehicles.These requirements include high-fidelity flow-fieldsolution methods, physical models, rapid designcapabilities and experimental techniques to developflowpath performance, flow-field characteristics,chemical kinetics, thermal management and aero-propulsive interactions. Specific research goals aredriven by systems analyses of candidatearchitectures and derived propulsion systemperformance goals. Current efforts and futuretechnology goals include mid-speed (Mach 3-8)combustion flow physics and componentdevelopment, hypervelocity (Mach 10-15) scramjetflow physics and flowpath performance, DMSJ andcombined-cycle flowpath performance andoperability, propulsion-airframe integration andcomputational and diagnostic tool development.Other enabling component technologies, such ashigh-temperature lightweight materials, sealsactuation mechanisms, and engine subsystems arealso part of the current program.

The paper will present a brief overview of

historical airframe-integrated scramjet propulsionresearch along with a brief discussion of currentresearch efforts within NASA. Technology shortfallsto develop Mach 3-15 scramjet propulsion flowpathsare discussed along with a proposed technologyroadmap to address these shortfalls and accomplishstrategic agency objectives.

Historical BackgroundFigure 2 shows a summary of scramjet

engine research at NASA’s Langley ResearchCenter. The first major scramjet engine developmentproject was the Hypersonic Research Engine (HRE),which began in the 1960’s.5 The goal of the HREproject was to flight test a flight-weight,regenera t i ve l y -coo led , hyd rogen- fue led ,axisymmetric scramjet engine on the X-15 researchairplane. The project was re-directed to ground testresearch when the X-15 program was terminated.Several component and flowpath tests wereconducted, including full-scale engine tests atNASA’s Langley and Lewis Research Centers. Thestructural assembly model (SAM) was tested in theLangley 8-Foot High Temperature Tunnel (8-Ft.HTT) and the Aerothermodynamic Integration Model(AIM) was tested in the Lewis Plumbrook HypersonicTest Facility (HTF).6,7 These tests demonstratedperformance and operability of a scramjet engine.

Following the HRE program, scramjet

Figure 2. Historical summary of NASA scramjet engine research.

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research within NASA was focused on airframe-integrated concepts. Research at NASA’s LangleyResearch Center focused on modular, fixed-geometry concepts, such as the 3-strut engine,utilizing sidewall compression. Sub-scale enginetests in the 1970’s demonstrated required thrust andoperability for an airframe-integrated scramjet-powered system. Two additional engine testprograms in the 1980’s included the strutlessparametric engine (SLPE) and the step-strutparametric engine (SSPE).

During the 1970’s and 1980’s, fundamentalresearch in scramjet engine design and performanceas well as supersonic combustion physics wasconducted at NASA Langley.8,9 These effortsincluded the development of empirical models formixing-controlled combustion, isolator performance,ignition and inlet operability limits. ComputationalFluid Dynamics (CFD) tools, including kineticsmodels for hydrogen-air combustion and otheraspects of supersonic chemically-reacting flowphysics modeling, were matured during this timeframe. Diagnostic tool development to obtaincalibration data for combustion models andcharacterize scramjet combustor flow physics wasalso conducted. Historical work in direct-connectcomponent tests to mature combustor design andmodeling tools is summarized in references 3 and 8.These include an investigation of a plasma torch asan ignition device source for hydrogen-fueledsupersonic combustors and investigations of variousfuel injector arrangements, including swept-ramp,expansion and compression ramps.

The National Aero-Space Plane (NASP) (X-30) program was initiated in the 1980’s to develop asingle-stage-to-orbit flight research vehicle. The X-30 concept proposed utilizing scramjet propulsion toMach 25. While the NASP program did not flight testan SSTO vehicle, major technology contributions toscramjet propulsion were accomplished. Theseincluded development of comprehensive Mach 3-8engine performance databases, hypervelocityscramjet performance and design methods, CFDand design tool maturation and propulsion-airframeintegration. Tests of the sub-scale parametric engine(SXPE) were conducted at NASA Langley in theArc-Heated Scramjet Test Facility (AHSTF).10 TheNASP Concept Demonstrator Engine (CDE), shownin figure 3, was tested in the Langley 8-Ft. HTT todemonstrate flowpath performance and operabilityand to verify flowpath design methods.11 During the1980’s, the 8-Ft. HTT underwent modifications toinstall a liquid oxygen (LOX) injection system toreplenish the oxygen consumed by the methane-air

combustion process.12 These modifications enabledtesting of large-scale hypersonic airbreathingpropulsion systems at flight enthalpies from Mach 4to 7. Comparisons of the SXPE and CDE test dataalso provided insight into ground test simulationconcerns, such as facility test gas composition,dynamic pressure and geometric scale effects.

NASA initiated the Hyper-X (X-43) flightresearch project in 1995 to demonstrate the in-flightperformance of a hydrogen-fueled, airframe-integrated scramjet at flight Mach numbers of 5, 7,and 10.13 The flight engine design was based on aMach 10, dual-fuel, global-reach reference vehicle.14

The project was subsequently redirected to focus onMach 7 and 10 flight tests, with ground engineresearch continuing at Mach 5.15 As part of theMach 7 flight engine and vehicle development,extensive ground f reejet engine andaerothermodynamic testing was conducted todemonstrate performance and operability and todevelop the engine and vehicle control laws.16,17

Figure 4 illustrates the X-43 ground engine test flowlogic. The dual-fuel experimental (DFX) engine wastested in the Langley AHSTF to characterizecombustor and flowpath performance. A hypersonicscramjet model (HSM) was tested in the NASAHYPULSE tunnel (reflected shock tunnel mode) toexamine the effects of pressure and facility test gasvitiation. The Hyper-X Engine Model (HXEM), a full-length, partial-width engine with truncated aftbody,was tested in the AHSTF and the 8-Ft. HTT. Thiscomparison also provided data on facility test gas

Figure 3. NASP Concept Demonstrator Engine.

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and dynamic pressure effects. The Hyper-X FlightEngine (HXFE), a full-scale duplicate flight engine,was tested extensively in the 8-Ft. HTT to verifyflight performance and control laws.18 The HXFE isshown in figure 5 during a run in the 8-Ft. HTT. Thefirst flight attempt of the Mach 7 X-43 in June 2000resulted in a failure of the booster rocket prior toreaching research vehicle separation and thescramjet test point. A second Mach 7 flight attempt isplanned in 2003.

Airbreathing launch vehicle and otherconcept studies continue to mature potentialconcepts for future flight demonstration andcandidate architectures for future operational

systems.19 These concept systems analyses areused to determine appropriate technologyinvestments and mature performance evaluations ofpotential vehicles. Studies have identified RBCC andTBCC propulsion concepts as potential candidatesfor SSTO systems. The TBCC concept integrates ahigh-speed turbojet and a scramjet in an“over/under” dual-flowpath configuration, depicted infigure 6.20

Current Research EffortsNASA’s Advanced Space Transportation

Program (ASTP) seeks to mature technologies toenable third-generation RLV systems.21 Thisprogram encompasses three aspects of hypersonicstechnology development: flight demonstration,ground demonstration and foundation technologyinvestments. Program implementation occurs atseveral NASA centers, including Marshall SpaceFlight Center, Langley Research Center, GlennResearch Center and Dryden Flight ResearchCenter.

Presently, the flight demonstration aspect of

Figure 5. Hyper-X Flight Engine (HXFE) in theLangley 8-Ft. High Temperature Tunnel (HTT). Figure 6. Over/Under Combination Cycle Concept.

Turbojet Scramjet

Figure 4. X-43 (Hyper-X) Ground Test Program Flow.

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ASTP consists of the X-43A (Hyper-X) and X-43Cprojects. The X-43C project is a joint NASA-AirForce project to achieve a flight demonstration of theUSAF HyTech engine, a hydrocarbon-fueled dual-mode scramjet (DMSJ).22 Similar to the X-43A, thismission is accomplished by boosting the researchvehicle to the flight test altitude and condition via asolid-rocket booster which is air-launched from a B-52 aircraft. Whereas the X-43A is designed toachieve only a few seconds of powered flight at asingle point design condition with heat-sinkhardware, the X-43C vehicle is designed to fly anaccelerating trajectory from Mach 5 to 7,demonstrating ramjet-scramjet mode transition.Additionally, the X-43C engine utilizes activeregenerative fuel cooling, which will provide avalidation of the heat exchanger design and theendothermic cooling capacity of liquid JP-7 fuel.

The two ground-based demonstrationprojects are formulated to enable development ofhydrocarbon-fueled combined-cycle propulsiontechnology up to Mach 7 conditions. The RBCCproject, led by NASA-Marshall Space Flight Center(MSFC), seeks to develop and ground test a Mach 7capable RBCC engine system.23-24 The proposedRBCC flowpath utilizes an ejector rocket system forapproximately Mach 0-3 operation, transitioning toramjet mode at approximately Mach 3 and finallytransitioning to scramjet mode for operation up toMach 7. The current RBCC ground demonstrationactivity builds upon earlier efforts to develophydrogen-fueled RBCC technology under theAdvanced Reusable Technologies (ART) program.25

The primary objective of the TBCC project, led byNASA-Glenn Research Center (GRC), is to develop,and demonstrate through ground testing, high-speedturbojet engines capable of operation up to Mach 4conditions. DMSJ performance and integration witha high-speed turbojet is also being studied at theconceptual level. A proposed X-43B sub-scalehydrocarbon-fueled reusable flight demonstratorvehicle is envisioned that would utilize either RBCCor TBCC propulsion for Mach 0.7 to 7 flight. Thisproject would extend the flight performancedatabase to low-speed systems, demonstrate modetransitions and address various operational aspectsof reusable flight vehicle technology.

These ground and flight demonstrationprojects support several DMSJ technologydevelopment areas. The X-43C project will matureregeneratively-cooled hydrocarbon-fueled scramjets.The RBCC project will examine rocket integration,mode transition, fueling strategies and enginecontrols. Inlet operability and fuel injector designs

have been examined in recently-conductedcomponent-level tests. The TBCC project willexamine over/under system integration andconceptual design of a DMSJ engine to functionover a Mach 4-7 flight trajectory.

The remaining foundation technologyaspects of the program are addressed in airframeand propulsion technology projects within ASTP.The airframe project includes maturation ofaerothermodynamics technologies as well aspropulsion-airframe integration and Mach 3-15integrated flowpath performance. Efforts in thepropulsion research project related to DMSJdevelopment include combustion flow physics andtool development as well as high-temperaturelightweight materials and seals. This paperdescribes current efforts and long-term researchobjectives in these foundation research andtechnology areas, including combustor technology,hypervelocity scramjet development, DMSJ andcontinuous performance and operability, propulsion-airframe integration and tool development.

Technology Development RoadmapTechnology development goals are driven

by the requirement for efficient DMSJ propulsionsystems for Mach 3-8 (near-term) and Mach 3-15(far-term) operation to support future flightdemonstration projects and operational space-access system development. Figure 7 shows asummary of current efforts as well as near-term andfar-term technology goals in each area.

Mid-Speed (Mach 3-8) Combustor TechnologyThe mid-speed flight regime generally refers

to the range from subsonic combustion operation tothe transition to supersonic operation. Generally, theterm is used to refer to the Mach 3-8 flight range.The combustor is characterized by highly distortedflow with regions of mixed subsonic and supersonicflow. In this context, the term “dual-mode” refers tothe region where mixed subsonic and supersonicflow is present in the combustor. The upstreampressure rise caused by heat release in thecombustor extends forward into the isolator section,which is characterized by an oblique shock train. Aprimary purpose of combustor component researchis to study the basic physical processes of fuelinjection and mixing as well operability of the isolatorand the combustor. Emphasis is placed on a betterunderstanding of low-speed (Mach 3-5)performance. In subsonic combustion mode, fuelinjection and combustion primarily occurdownstream in the diverging section of thecombustor and the heat release due to combustion

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creates a thermal throat. The primary challenge incombustor design is determining appropriate fuelingstrategies and injector designs to improveperformance in and through this rapid expansionregion. Parametric combustor design databases areneeded in this speed regime to characterizeperformance and improve multi-speed combustoroptimization.

An understanding of combustor operationand the design application to future hypersonic flightdemonstrators and operational vision vehicles isfurther complicated by fuel selection. Hydrogen fuelis the preferred option for single-stage-to-orbitairbreathing launch vehicles with scramjet cycles toMach 15 or greater. However, liquid hydrocarbonfuels have beneficial applications because of theirgreater fuel densities and endothermic coolingcapabilities. There is a generally accepted upperlimit of approximately Mach 8 for storable JP-typehydrocarbon fuels.26,27 An advantage of liquidhydrocarbon fuels for the proposed combined-cycleflight demonstrators is smaller vehicle designs andtherefore, potential cost savings. Also, the use ofhydrocarbon fuels also extends the applicability oftechnology development to hypersonic cruisemissions. In addition to the hydrocarbon-fueledcombustor development in the ground and flightdemonstration projects previously discussed, thereis a need for fundamental supersonic combustionstudies for hydrogen-fueled engines, including an

understanding of performance relationships andscaling parameters between hydrogen andhydrocarbon-fueled combustor designs.

The Langley Direct-Connect SupersonicCombustion Test Facility (DCSCTF), depicted infigure 8, is devoted to DMSJ combustor testing anddevelopment.10 This facility utilizes hydrogen-aircombustion to achieve a test gas that duplicates thestagnation enthalpies of flight Mach numbers from4.0 to 7.5. Oxygen replenishment is used to achievea test gas with the same oxygen mole fraction asatmospheric air (0.2095). Gaseous hydrogen is theprincipal fuel used in combustor models tested in theDCSCTF, although other gas mixtures, such asethylene, have been used to simulate crackedhydrocarbon fuels. Modifications are in progress toaccommodate testing with heated liquid hydrocarbonfuels. Near-term efforts will consist of parametricdirect-connect combustor testing with both gaseousand liquid fuels. A parametric combustor test articleis proposed which will allow for the investigation ofcombustor design parameters, in jectorconfigurations, fueling strategies, ignition andflameholding characteristics. These efforts willprovide risk reduction for future flight demonstratorengine development, specifically the RBCC andTBCC engine concepts and will generate aparametric design database for multi-speedcombustor optimization. These efforts are coupledwith investigations of combustion flow physics,

Hypervelocity (Mach 10-15) Scramjet Engine Development

Mid-Speed (Mach 3-8) Combustor Technology

Mach 3-8 DMSJ/Combined-Cycle Performance and Operability

Propulsion-Airframe Integration

Analysis Tools and Diagnostics Development

Near-Term Foci

• Hypervelocity Scramjet/RBCCGround Demonstration

• Parametric Combustor Test Rig• M 3-15 multi-speed optimization

• Variable-Geometry PerformanceDemonstration

• Multi-speed Aero-PropulsivePerformance Methods

• Physical and Algorithmicenhancements and validation

• Temperature, species and velocitymeasurements.

Current Efforts

• Flow Physics and Test Technques

• Flow Physics and Modeling

• Mach 3-6 Performance Data• Engine Control Algorithms

• Powered Transonic Aero• Sub-scale Simulation

• Code Development & Validation• Combustor Flowfield Temperature

Measurements

Far-Term Goals

• LOx-augmented Ejector-ScramjetDemonstration

• Flight-weight H2-FueledCombustor Validation.

• H2-Fueled Flight-like, Fuel-Cooled Engine Demonstration

• Aero-Propulsive TrajectorySimulation Capability

• Rapid nose-to-tail design tools• Routine measurements of

complete combustor flowfieldproperties

Figure 7. Hydrogen-Fueled DMSJ Technology Roadmap

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computational modeling and associated diagnostictool development (described in the next section).Far-term goals are to design and demonstrate aflight-weight, regeneratively-cooled, hydrogen-fueledcombustor for future Mach 0-15 flight demonstrationand vision vehicle development.

Analysis Tool and Diagnostic DevelopmentComputational Fluid Dynamics (CFD)

modeling of scramjet combustors is complicated byvarious physical processes, including large regionsof subsonic flow, separated flow regions, complexmixing phenomena, non-equilibrium transfer ofturbulence energy, and interactions betweenturbulence and chemical kinetics that may impactboth the chemical reactions and turbulence field.26

Limitations in physical modeling capabilities as wellas computational overhead costs limit the practicaluse of three-dimensional CFD tools in design anddevelopment of scramjet combustors. Several effortsare underway to incorporate advanced physicalmodels and algorithmic enhancements in CFD toolsand to acquire high-fidelity data sets for codecalibration.

The Coherent anti-Stokes RamanSpectroscopy (CARS) technique has been usedrecently to acquire flow-field data in the LaRCDCSCTF for this purpose.28,29 These efforts havesucceeded in mapping mean and RMS temperaturefluctuations in a supersonic combustor. Figure 9shows a generic supersonic combustor model withangular fuel injection used for these studies andfigure 10 shows mean temperature maps developed

using the CARS technique.28 The experiment wasadopted as a test case by the RTO working groupon scramjet propulsion (Working Group 10).29 Near-term objectives of this work include furtherapplications of optical diagnostic techniques toobtain simultaneous temperature and speciesmeasurements. The VULCAN CFD code, developedat NASA LaRC, is a Navier-Stokes code used tosimulate chemically-reacting flow fields in scramjetcombustors and flowpaths.31 Previously, upgrades tothis code, including volume grid capabilities, physicalmodels and convergence accelerat ionenhancements were incorporated. The combustordata sets described here comprise a partial database for CFD code validation for this class of flows.

Additional work has included the use oflaser-based diagnostics along with flow seeding toobtain fuel plume images and velocitymeasurements. Reference 31 described the use ofthis technique to obtain measurements in a Mach 2hydrogen-air combustor with a 10o unswept rampfuel injector. The VULCAN CFD code was used toobtain flow field predictions of this flowpath andcomparisons with experimental wall pressures andfuel plume images are shown. These comparisonsindicate an underprediction in the level of turbulentmixing and heat release due to combustion. Thiswork suggests that improvements in turbulencemodeling and turbulent-chemistry interactions areneeded to improve supersonic combustor modelingin this speed range.

Future efforts are expected to focus on thedevelopment of validated nose-to-tail design andanalysis tools. This includes advanced physicalmodeling capabilities for turbulence, turbulence-chemistry interactions and reduced combustionkinetics models for hydrocarbon and hydrogencombustion. Ultimately, diagnostic tool developmentwill enable routine measurements of all combustor

flow field parameters (temperature, velocity, species

Combustor Model

Facility Nozzle

Vitiated Air Stream

Divergent Section

Figure 8. Langley Direct-Connect SupersonicCombustion Test Facility (DCSCTF).

Figure 9. Supersonic Combustor (SCHOLAR)Model.

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concentrations) to provide the necessary databasesto fully characterize combustor flow fields and tovalidate CFD methods.

DMSJ/Combined-Cycle Performance and OperabilityDual-mode scramjet engine cycles must

function over a wide Mach number range in order tomeet mission requirements. Current ground testdatabases exist only over a limited Mach numberrange. For candidate X-43B demonstratorconfigurations, engine performance, operability andsystem weights must enable accomplishment of theMach 7 mission objective. Future Mach 15 andSSTO operational concepts require scramjetoperation up to Mach 15 before transitioning to arocket cycle or scramjet ejector cycle for orbitalinsertion. Therefore, mission specific impulse valuesmust be maximized over the Mach number range inorder to provide acceptable performance margin.

Historically, a number of engine testprograms at NASA-Langley, discussed previously,have contributed to the performance database forair frame-integrated scramjet f lowpaths.3,4

Depending on the Mach number range and specificmission requirements, efficient inlet operation overthe applicable flight regime may necessitatevariable contraction ratios. Variable-geometryconcepts have been examined to provide thisrequired operability.

Efficient multi-speed engine operation alsorequires the development of fueling strategies and

engine control mechanisms. Engine controlalgorithms are required to control the enginevariable contraction ratio schedule and control fuelflow rates to provide required thrust to meet missionobjectives, enable mode transitions and to preventand recover from engine unstart and flameout.These issues were studied in the Hyper-X programfor the Mach 7 and 10 flight experiments andcontributed to the development of flight control lawsfor the X-43A vehicles. This research is being furtherextended to the Mach 4-7 range and matured forfuture flight demonstrator vehicle development.

Near-term project plans consist of additionalfreejet testing efforts in the Langley CHSTF and theAHSTF. First, the existing Mach 5 dual-fuelexperimental engine (DFX) will be tested in theCHSTF to extend the low-speed (Mach 3-5)hydrogen-fueled engine performance database forDMSJ engines and to further examine facility test-gas vitiation effects. Second, further testing of theHXEM will be conducted in the AHSTF to investigatecontrol laws for scramjet engine operation over theMach 4-7 range. The HXEM testing will take placefollowing the installation of a new, high-quality flow,Mach 6 facility nozzle in the tunnel circuit. Athorough calibration at three total enthalpy levels willbe conducted followed by HXEM performancetesting to examine the effects of nozzle exit flowuniformity on engine performance.

A longer-term objective of the program is toconduct a comprehensive test program of a variable-geometry dual-mode scramjet engine configurationover the Mach 3-8 speed range. The goal of this testprogram will be to investigate parametricperformance at fixed contraction ratios, examinecontraction ratio changes during runs, examineperformance during mode transitions and to verifyclosed-loop engine control algorithms. Real-timeenthalpy and dynamic pressure variation duringtests will be accomplished. Variable-geometrydemonstration with heat-sink hardware will befollowed with flight-like, regeneratively-cooled,hydrogen-fueled engine ground testing.

Hypervelocity Scramjet DevelopmentAs indicated in figure 1, the specific impulse

of the scramjet cycle decreases as Mach numberincreases. Heat release due to combustion isinversely proportional to the square of thefreestream Mach number. At Mach 15, thecombustion energy is approximately less than 25-percent of the free stream kinetic energy, accountingfor flow field losses.9 At these Mach numbers, smallchanges in effective specific impulse can cause

Figure 10. Mean temperature maps obtained inthe DCSCTF using CARS (ref. 25).

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significant changes in vehicle take-off gross weights,thus impacting the ability of the system to meetmission performance requirements.3 2 Thisrepresents a practical upper limit for efficientscramjet engine operation without LOX-augmentation. A further understanding of thefundamental physical processes that govern engineperformance in the hypervelocity speed range isrequired in order to optimize flowpath lines forefficient operation.

At hypervelocity speeds, scramjet flowphysics are characterized by very short residencetimes. Robust flowpath design is dependent on anunderstanding of the following flow-field andcombustion phenomena: fuel injector geometry,mixing, flameholding, and combustion efficiency aswell as thermal balance and protectionrequirements. The NASA-Langley HYPULSE facility,located at and operated by Allied Aerospace, Inc.(GASL Division) has been used to performhypervelocity scramjet research in a shock-expansion-tunnel (SET) mode at enthalpy levelsduplicating Mach numbers above 10.33-35 Aschematic of the NASA HYPULSE Facility is shownin figure 11. Tests have been conducted mostrecently on a scramjet flowpath model,representative of the Mach 10 X-43A scramjetflowpath lines, at Mach 15 conditions.36

Future test technique development will focuson a definition and calibration of HYPULSE (SET) atbaseline test points in the Mach 12-15 range. Thisincludes the design, using CFD, and fabrication of afacility nozzle suitable for scramjet engine tests, andefforts to optimize and calibrate the shock tunnel exitflow conditions for these flight Mach numbers.Various flow diagnostic techniques, includingschlieren, fuel-plume (planar) imaging, watertemperature and concentration by laser absorptionand laser holographic interferometry, will be applied

to assess facility test conditions.

Near-term efforts also consist ofdevelopment of a comprehensive Mach 12-15performance database. This will enable the designof a future flight demonstrator to validate Mach 15scramjet performance. Long-term goals includedemonstrations of a flight-weight combustor at Mach15 conditions and LOX-augmented ejector-scramjetcycles to enable orbital insertion for SSTO systems.System studies indicate that LOX-augmentation maybe required for efficient orbital insertion in SSTOairbreathing launch vehicles.37

Propulsion-Airframe Integration (PAI)Hypersonic airbreathing vehicles are

characterized by highly integrated systems with ahigh degree of interaction between the airframe andp r o p u l s i o n flowpath. Aerothermodynamicperformance cannot be decoupled from propulsionperformance, due to shared surfaces and flow fieldinteractions, as depicted in figure 12. Therefore, asignificant challenge in the design and developmentof hypersonic airbreathing flight vehicles is thedetermination of aero-propulsive interactions andinstalled vehicle performance. Design tools andground testing techniques are needed to fullycharacterize these effects across the applicablespeed ranges.

During the NASP program, significant workwas done to investigate the use of cold-gas mixturesto simulate powered scramjet exhaust products inground test facilities.38-40 This technique wasinvestigated in the supersonic and hypersonic speedregimes (Mach 4-10) with powered metric aftbodymodels to develop the technique and measureexternal nozzle pressures, exhaust plumeimpingement on wing surfaces and aftbody forcesand moments. Analysis to examine the correlation ofcold simulant gases to hot combustion products wasinitiated, but not completed due to the termination ofthis program.

The X-43A flight project undertook asignificant ground testing and computational effort tobuild the pre-flight vehicle database.41,42 This effortconsisted of un-powered aerothermodynamic testingwith powered force and moment incrementssupplied by CFD predictions. These predictions wereverified by full-flowpath force and moment dataobtained from the HXFE testing in the Langley 8-Ft.HTT.18

Future NASA efforts in this area will build onNASP and X-43A research and will focus on testFigure 11. NASA HYPULSE Facility.

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techniques and analysis methodologies to predictinstalled vehicle forces and moments through theappl icab le f l ight t ra jec tor ies . In thesupersonic/hypersonic speed range, data from sub-scale exhaust simulation methodologies will becompared with data from full-flowpath fueledscramjet test models to further mature this techniqueand quantify the uncertainty in force and momentmeasurements.

Vehicle systems studies have identifiedinstalled transonic performance predictions ashaving a high degree of uncertainty. Currentanalytical methods are not sufficient to fullycharacterize the integrated powered vehicleperformance in this speed range. Recently, a testwas conducted in the Langley 16-Foot TransonicTunnel to measure nozzle/aftbody flow fieldcharacteristics and forces and moments on a NASPvehicle model. These data will be used to calibrateexisting CFD and other design methodologies forfuture analyses.

Additionally, maturation of test techniques tomeasure vehicle forces and moments andpropulsion-airframe interactions are required for thehypervelocity (Mach 10-15) speed range. Typically,pulse facilities are used to supply the energy neededfor aerothermodynamic and propulsion testing in thisspeed range. Measurement of powered forces andmoments in these environments is a technique notadequately addressed in previous NASA hypersonicprograms. The University of Queensland haspreviously reported efforts to support aerodynamicforce and moment measurements in pulse facilities.

Reference 43 reports three-componentmeasurements obtained in a reflected shock tunneland reference 44 describes a 6-component balancedesign. These devices rely on a stress wave forcemeasurement technique, which measures thepropagation of stress waves through test models.

The TBCC over/under configuration requiresa high degree of integration to determine the impacton the performance of each flowpath and thetransition regime from gas turbine to scramjet. Dual-flowpath inlet and nozzle integration test articles areenvisioned to demonstrate the integration and modetransition aspects of this system.

Other aerothermodynamic technologies thatimpact propulsion system operation are also beingaddressed in other areas of the ASTP airframetechnology project. These include boundary-layertransition studies, evaluation of alternative forcedtransition mechanisms, shock interaction methodsand aeroheating prediction methods. These effortsare fully described in reference 45.

Other Enabling TechnologiesIn addition to the scramjet combustor and

flowpath development efforts described herein,additional key enabling technologies will be requiredto accomplish operational system goals. Theseinclude lightweight high-temperature materials andseals, actuation mechanisms, hot-gas valves andfuel systems. Technology development in theseareas is either captured in existing projects or theywill be addressed in future focused flight enginedevelopment.

Concluding RemarksThe U.S. National Aeronautics and Space

Administration (NASA) is developing technologies toenable third-generation reusable launch vehicles(RLV). The focus is on airbreathing and combined-cycle propulsion concepts to take advantage ofpotential performance gains over conventionalrocket propulsion. The dual-mode scramjet (DMSJ)engine cycle, integrated with rocket or turbine-basedcycles, is being studied for future RLV vision vehicleconcepts. Historically, development of airframe-integrated scramjets has progressed throughcomponent testing and sub-scale engine testprograms, the National Aerospace Plane (NASP)program and the current X-43A and X-43C flightdemonstration projects. Foundation technologyinvestments, supporting propulsion grounddemonstration projects, future flight demonstratorsand vision vehicle operational system concepts arerequired to achieve strategic goals. A proposed

Exhaust PlumeInteraction with

Airframe,Control Surfaces

Exhaust Plume andNozzle Expansion

Impact VehiclePitching Moments

Forebody Shaping, Shock Structure andBoundary-Layer State Impacts Engine Mass

Capture, Combustion Efficiency

Variable-GeometryEngine Operation

Figure 12. Propulsion-Airframe Integrationconsiderations for hypersonic vehicles.

AIAA 2002-5188


NASA technology roadmap has been presentedwhich encompasses mid-speed (Mach 3-8) dual-mode combustion flow physics, hypervelocity (Mach10-15) scramjet development, multi-speed flowpathoperability and performance, propulsion-airframeintegration and analysis and diagnostic tooldevelopment.

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Documents

AIAA 2002-5188 Technology Roadmap for Dual-Mode …mln/ltrs-pdfs/NASA-aiaa-2002-5188.pdf · AIAA 2002-5188 1 AIAA/AAAF ... fundamental research in supersonic combustion flow physics,