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NASA Technical Reports Server (NTRS) 20030000444: Technology Roadmap for Dual-Mode Scramjet Propulsion to Support Space-Access Vision Vehicle Development PDF

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American Institute of Aeronautics and Astronautics AIAA 2002-5188 Technology Roadmap for Dual-Mode Scramjet Propulsion to Support Space-Access Vision Vehicle Development Charles E. Cockrell, Jr. Aaron H. Auslender R. Wayne Guy Charles R. McClinton Sharon S. Welch NASA Langley Research Center Hampton, Virginia, USA 11th AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference 29 September- 4 October 2002 / Orleans, France 1 AIAA 2002-5188 Technology Roadmap for Dual-Mode Scram jet Propulsion to Support Space-Access Vision Vehicle Development Charles E. Cockrell, Jr.*, Aaron H.Auslender t, R.Wayne GuyS,Charles R. McClinton §,Sharon S. Welch §§ NASA Langley Research Center, Hampton, VA to-space roadmap has been established that Abstract focuses on airframe-integrated hypersonic airbreathing propulsion development through Third-generation reusable launch vehicle foundation technology investments, ground (IRLV) systems are envisioned that utilize demonstration and flight validation. Successful airbreathing and combined-cycle propulsion to take implementation of this roadmap requires a robust advantage of potential performance benefits over technology development program to mature aspects conventional rocket propulsion and address goals of of the propulsion system and integrated aero- reducing the cost and enhancing the safety of propulsive vehicle performance through both systems to reach earth orbit. The dual-mode analytic and experimental research. scramjet (DMSJ) forms the core of combined-cycle or combination-cycle propulsion systems for single- Figure 1 shows a comparison of nominal stage-to-orbit (SSTO) vehicles and provides most of specific-impulse values for airbreathing engine the orbital ascent energy. These concepts are also cycles vs. rockets. The dual-mode scramjet (DMSJ) relevant to two-stage-to-orbit (TSTO) systems with forms the core of combined-cycle or combination- an airbreathing first or second stage. Foundation cycle airbreathing propulsion systems and provides technology investments in scramjet propulsion are most of the orbital ascent energy for single-stage-to- driven by the goal to develop efficient Mach 3-15 orbit (SSTO) airbreathing launch vehicle systems. concepts with sufficient performance and operability The term "dual-mode scramjet" refers to an engine to meet operational system goals. A brief historical cycle that can operate in both subsonic combustion review of NASA scramjet development is presented and supersonic combustion modes. Rocket-based along with a summary of current technology efforts combined cycle (IRBCC) concepts are being studied and a proposed roadmap. The technology which integrate rocket thrusters with the DMSJ addresses hydrogen-fueled combustor development, flowpath for low-speed propulsion. Turbine-based hypervelocity scramjets, multi-speed flowpath combination cycle (TBCC) concepts are also being performance and operability, propulsion-airframe examined which integrate a gas turbine engine and integration, and analysis and diagnostic tools. DMSJ in a dual-flowpath configuration. Introduction Hypersonic airbreathing propulsion research The United States National Aeronautics and conducted by NASA spans over 40 years. 1-4 Space Administration (NASA) has established a Historical work includes the hypersonic research strategic goal of creating a safe, affordable highway engine (HIRE), airframe-integrated scramjet ground through the air and into space. Candidate third- testing and component development, the X-30 generation reusable launch vehicle (RLV) National Aerospace Plane (NASP) program, and, architectures include single-stage and two-stage more recently, the Hyper-X (X-43) flight concepts which utilize airbreathing, combined-cycle and combination-cycle propulsion systems to take advantage of potential performance gains over conventional rocket-propelled concepts. An access- Isp * AirbreathingPropulsionProgramManager,AdvancedSpace TransportationProgramOffice,SeniorMember,AIAA. t AssistantHead,HypersonicAirbreathingPropulsionBranch. :1:Head,HypersonicAirbreathingPropulsionBranch. § TechnologyManager,Hyper-XProgramOffice. §§Head,AdvancedSpaceTransportationProgramOffice. Copyright © 2002 by American Institute of Aeronautics and Astronautics,Inc.NocopyrightisassertedintheUnitedStates underTitle17,U.S.Code.TheU.S.Governmenthasa royalty- free license to exerciseall rights underthe copyrightclaimed 10 20 Mach Number hereinforgovernmentalpurposes.All otherrightsare reserved bythecopyrightowner. Figure 1. Airbreathing Propulsion Performance. AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference AIAA2002-5188 demonstration project. This work has also included historical airframe-integrated scramjet propulsion fundamental research in supersonic combustion flow research along with a brief discussion of current physics, analysis tools and diagnostic research efforts within NASA. Technology shortfalls methodologies. to develop Mach 3-15 scramjet propulsion flowpaths are discussed along with a proposed technology Multidisciplinary research is required to roadmap to address these shortfalls and accomplish develop efficient airbreathing and combined-cycle strategic agency objectives. propulsion systems for airframe-integrated vehicles. These requirements include high-fidelity flow-field Historical Backqround solution methods, physical models, rapid design Figure 2 shows a summary of scramjet capabilities and experimental techniques to develop engine research at NASA's Langley Research flowpath performance, flow-field characteristics, Center. The first major scramjet engine development chemical kinetics, thermal management and aero- project was the Hypersonic Research Engine (HRE), propulsive interactions. Specific research goals are which began in the 1960's. 5 The goal of the HRE driven by systems analyses of candidate project was to flight test a flight-weight, architectures and derived propulsion system regeneratively-cooled, hydrogen-fueled, performance goals. Current efforts and future axisymmetric scramjet engine on the X-15 research technology goals include mid-speed (Mach 3-8) airplane. The project was re-directed to ground test combustion flow physics and component research when the X-15 program was terminated. development, hypervelocity (Mach 10-15) scramjet Several component and flowpath tests were flow physics and flowpath performance, DMSJ and conducted, including full-scale engine tests at combined-cycle flowpath performance and NASA's Langley and Lewis Research Centers. The operability, propulsion-airframe integration and structural assembly model (SAM) was tested in the computational and diagnostic tool development. Langley 8-Foot High Temperature Tunnel (8-Ft. Other enabling component technologies, such as HTT) and the Aerothermodynamic Integration Model high-temperature lightweight materials, seals (AIM) was tested in the Lewis Plumbrook Hypersonic actuation mechanisms, and engine subsystems are Test Facility (HTF). 6'7 These tests demonstrated also part of the current program. performance and operability of a scramjet engine. The paper will present a brief overview of Following the HRE program, scramjet Figure 2. Historical summary of NASA scram jet engine research. AIAA 2002-5188 research within NASA was focused on airframe- combustion process. 12These modifications enabled integrated concepts. Research at NASA's Langley testing of large-scale hypersonic airbreathing Research Center focused on modular, fixed- propulsion systems at flight enthalpies from Mach 4 geometry concepts, such as the 3-strut engine, to 7. Comparisons of the SXPE and CDE test data utilizing sidewall compression. Sub-scale engine also provided insight into ground test simulation tests in the 1970's demonstrated required thrust and concerns, such as facility test gas composition, operability for an airframe-integrated scramjet- dynamic pressure and geometric scale effects. powered system. Two additional engine test programs in the 1980's included the strutless NASA initiated the Hyper-X (X-43) flight parametric engine (SLPE) and the step-strut research project in 1995 to demonstrate the in-flight parametric engine (SSPE). performance of a hydrogen-fueled, airframe- integrated scramjet at flight Mach numbers of 5, 7, During the 1970's and 1980's, fundamental and 10.13 The flight engine design was based on a research in scramjet engine design and performance Mach 10, dual-fuel, global-reach reference vehicle. 14 as well as supersonic combustion physics was The project was subsequently redirected to focus on conducted at NASA Langley. 8'9 These efforts Mach 7 and 10 flight tests, with ground engine included the development of empirical models for research continuing at Mach 5.15 As part of the mixing-controlled combustion, isolator performance, Mach 7 flight engine and vehicle development, ignition and inlet operability limits. Computational extensive ground freejet engine and Fluid Dynamics (CFD) tools, including kinetics aerothermodynamic testing was conducted to models for hydrogen-air combustion and other demonstrate performance and operability and to aspects of supersonic chemically-reacting flow develop the engine and vehicle control laws. 16'17 physics modeling, were matured during this time Figure 4 illustrates the X-43 ground engine test flow frame. Diagnostic tool development to obtain logic. The dual-fuel experimental (DFX) engine was calibration data for combustion models and tested in the Langley AHSTF to characterize characterize scram jet combustor flow physics was combustor and flowpath performance. A hypersonic also conducted. Historical work in direct-connect scramjet model (HSM) was tested in the NASA component tests to mature combustor design and HYPULSE tunnel (reflected shock tunnel mode) to modeling tools is summarized in references 3 and 8. examine the effects of pressure and facility test gas These include an investigation of a plasma torch as vitiation. The Hyper-X Engine Model (HXEM), a full- an ignition device source for hydrogen-fueled length, partial-width engine with truncated aftbody, supersonic combustors and investigations of various was tested in the AHSTF and the 8-Ft. HTT. This fuel injector arrangements, including swept-ramp, comparison also provided data on facility test gas expansion and compression ramps. The National Aero-Space Plane (NASP) (X- 30) program was initiated in the 1980's to develop a single-stage-to-orbit flight research vehicle. The X- 30 concept proposed utilizing scramjet propulsion to Mach 25. While the NASP program did not flight test an SSTO vehicle, major technology contributions to scramjet propulsion were accomplished. These included development of comprehensive Mach 3-8 engine performance databases, hypervelocity scramjet performance and design methods, CFD and design tool maturation and propulsion-airframe integration. Tests of the sub-scale parametric engine (SXPE) were conducted at NASA Langley in the Arc-Heated Scramjet Test Facility (AHSTF). 1° The NASP Concept Demonstrator Engine (CDE), shown in figure 3, was tested in the Langley 8-Ft. HTT to demonstrate flowpath performance and operability and to verify flowpath design methods. 11During the 1980's, the 8-Ft. HTT underwent modifications to install a liquid oxygen (LOX) injection system to replenish the oxygen consumed by the methane-air Figure 3. NASP Concept Demonstrator Engine. 4 AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference AIAA2002-5188 MACH 7 HYPER-X ENGINE TEST PROGRAM _j_::'_* ......\..."';',:',_!,:\_"6])------_" _,_J_ ................................:.:.:.:.:.:.:.:.:.:.:.:.:.::::::::::::::::: ........._...:::::::[: ....................................-. \ / ]. GASL HYPULSE T_"d_Co-3&}d F_, _"}.mC*t*,_d F.B. fE..JI .......... HXEM p'__ ...................==========================.=..=..==================================== _'_':'_1._,,,t"_"_ T_,_,_ F_. Clean atr Archeated CH4-O2-Afr Figure 4. X-43 (Hyper-X) Ground Test Program Flow. and dynamic pressure effects. The Hyper-X Flight systems. These concept systems analyses are TM Engine (HXFE), a full-scale duplicate flight engine, used to determine appropriate technology was tested extensively in the 8-Ft. HTT to verify investments and mature performance evaluations of flight performance and control laws. _8The HXFE is potential vehicles. Studies have identified RBCC and shown in figure 5 during a run in the 8-Ft. HTT. The TBCC propulsion concepts as potential candidates first flight attempt of the Mach 7 X-43 in June 2000 for SSTO systems. The TBCC concept integrates a resulted in a failure of the booster rocket prior to high-speed turbojet and a scramjet in an reaching research vehicle separation and the "over/under" dual-flowpath configuration, depicted in scramjet test point. A second Mach 7flight attempt is figure 6.20 planned in 2003. Current Research Efforts Airbreathing launch vehicle and other NASA's Advanced Space Transportation concept studies continue to mature potential Program (ASTP) seeks to mature technologies to concepts for future flight demonstration and enable third-generation RLV systems. 2_ This candidate architectures for future operational program encompasses three aspects of hypersonics technology development: flight demonstration, ground demonstration and foundation technology investments. Program implementation occurs at several NASA centers, including Marshall Space Flight Center, Langley Research Center, Glenn Research Center and Dryden Flight Research Center. Presently, the flight demonstration aspect of Turbojet Scram jet ..... _'_,:_ Figure 5. Hyper-X Flight Engine (HXFE) inthe Langley 8-Ft. High Temperature Tunnel (HTT). 5 Figure 6. Over/Under Combination Cycle Concept. AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference AIAA 2002-5188 ASTP consists of the X-43A (Hyper-X) and X-43C have been examined in recently-conducted projects. The X-43C project is a joint NASA-Air component-level tests. The TBCC project will Force project to achieve a flight demonstration of the examine over/under system integration and conceptual design of a DMSJ engine to function USAF HyTech engine, a2hydrocarbon-fueled dual- mode scramjet (DMSJ). Similar to the X-43A, this over a Mach 4-7 flight trajectory. mission is accomplished by boosting the research vehicle to the flight test altitude and condition via a The remaining foundation technology solid-rocket booster which is air-launched from a B- aspects of the program are addressed in airframe 52 aircraft. Whereas the X-43A is designed to and propulsion technology projects within ASTP. achieve only a few seconds of powered flight at a The airframe project includes maturation of single point design condition with heat-sink aerothermodynamics technologies as well as hardware, the X-43C vehicle is designed to fly an propulsion-airframe integration and Mach 3-15 accelerating trajectory from Mach 5 to 7, integrated flowpath performance. Efforts in the demonstrating ramjet-scramjet mode transition. propulsion research project related to DMSJ Additionally, the X-43C engine utilizes active development include combustion flow physics and regenerative fuel cooling, which will provide a tool development as well as high-temperature validation of the heat exchanger design and the lightweight materials and seals. This paper endothermic cooling capacity of liquid JP-7 fuel. describes current efforts and long-term research objectives in these foundation research and The two ground-based demonstration technology areas, including combustor technology, projects are formulated to enable development of hypervelocity scramjet development, DMSJ and hydrocarbon-fueled combined-cycle propulsion continuous performance and operability, propulsion- technology up to Mach 7 conditions. The RBCC airframe integration and tool development. project, led by NASA-Marshall Space Flight Center (MSFC), seeks to develop and ground test a Mach 7 Technology Development Roadmap 23 24 capable RBCC engine system. - The proposed Technology development goals are driven RBCC flowpath utilizes an ejector rocket system for by the requirement for efficient DMSJ propulsion approximately Mach 0-3 operation, transitioning to systems for Mach 3-8 (near-term) and Mach 3-15 ramjet mode at approximately Mach 3 and finally (far-term) operation to support future flight transitioning to scramjet mode for operation up to demonstration projects and operational space- Mach 7. The current RBCC ground demonstration access system development. Figure 7 shows a activity builds upon earlier efforts to develop summary of current efforts as well as near-term and hydrogen-fueled RBCC technology under the far-term technology goals ineach area. Advanced Reusable Technologies (ART) program. 25 The primary objective of the TBCC project, led by Mid-Speed (Mach 3-8) Combustor Technoloqy NASA-Glenn Research Center (GRC), is to develop, The mid-speed flight regime generally refers and demonstrate through ground testing, high-speed to the range from subsonic combustion operation to turbojet engines capable of operation up to Mach 4 the transition to supersonic operation. Generally, the conditions. DMSJ performance and integration with term is used to refer to the Mach 3-8 flight range. a high-speed turbojet is also being studied at the The combustor is characterized by highly distorted conceptual level. A proposed X-43B sub-scale flow with regions of mixed subsonic and supersonic hydrocarbon-fueled reusable flight demonstrator flow. In this context, the term "dual-mode" refers to vehicle is envisioned that would utilize either RBCC the region where mixed subsonic and supersonic or TBCC propulsion for Mach 0.7 to 7 flight. This flow is present in the combustor. The upstream project would extend the flight performance pressure rise caused by heat release in the database to low-speed systems, demonstrate mode combustor extends forward into the isolator section, transitions and address various operational aspects which is characterized by an oblique shock train. A of reusable flight vehicle technology. primary purpose of combustor component research is to study the basic physical processes of fuel These ground and flight demonstration injection and mixing as well operability of the isolator projects support several DMSJ technology and the combustor. Emphasis is placed on a better development areas. The X-43C project will mature understanding of low-speed (Mach 3-5) regeneratively-cooled hydrocarbon-fueled scram jets. performance. In subsonic combustion mode, fuel The RBCC project will examine rocket integration, injection and combustion primarily occur mode transition, fueling strategies and engine downstream in the diverging section of the controls. Inlet operability and fuel injector designs combustor and the heat release due to combustion 6 AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference AIAA2002-5188 CurrentEfforts Near-TermFoci Far-TermGoals iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii ii iii iiiii i i i ii iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii i iii iii iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii iiiiiiiiiiiiiiiiii_iiiiiiiiiiiiiiiiiiiiiiiiiii_i_iiiiiii_ii_iiiii_i_iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii__iii_i_iii_i_ii_i_iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii_iiiiiiiiiiiiiiiiii_i_iii_i_i_iii_!iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii iiiiiiiiiiiiiiiiiiiiiiiiM i iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii iiiiiiiiiiiiiiiii iiii i i iii i iii i iiiiiiiiiiiiiiiiiiiii iii i ii diiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii ii`i i `ii`iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii i i"i iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii Figure 7. Hydrogen-Fueled DMSJ Technology Roadmap creates a thermal throat. The primary challenge in understanding of performance relationships and combustor design is determining appropriate fueling scaling parameters between hydrogen and strategies and injector designs to improve hydrocarbon-fueled combustor designs. performance in and through this rapid expansion region. Parametric combustor design databases are The Langley Direct-Connect Supersonic needed in this speed regime to characterize Combustion Test Facility (DCSCTF), depicted in performance and improve multi-speed combustor figure 8, is devoted to DMSJ combustor testing and optimization. development. _° This facility utilizes hydrogen-air combustion to achieve a test gas that duplicates the An understanding of combustor operation stagnation enthalpies of flight Mach numbers from and the design application to future hypersonic flight 4.0 to 7.5. Oxygen replenishment is used to achieve demonstrators and operational vision vehicles is a test gas with the same oxygen mole fraction as further complicated by fuel selection. Hydrogen fuel atmospheric air (0.2095). Gaseous hydrogen is the is the preferred option for single-stage-to-orbit principal fuel used in combustor models tested inthe airbreathing launch vehicles with scramjet cycles to DCSCTF, although other gas mixtures, such as Mach 15 or greater. However, liquid hydrocarbon ethylene, have been used to simulate cracked fuels have beneficial applications because of their hydrocarbon fuels. Modifications are in progress to greater fuel densities and endothermic cooling accommodate testing with heated liquid hydrocarbon capabilities. There is a generally accepted upper fuels. Near-term efforts will consist of parametric limit of approximately Mach 8 for storable JP-type direct-connect combustor testing with both gaseous hydrocarbon fuels. 26 '27 An advantage of liquid and liquid fuels. A parametric combustor test article hydrocarbon fuels for the proposed combined-cycle is proposed which will allow for the investigation of flight demonstrators is smaller vehicle designs and combustor design parameters, injector therefore, potential cost savings. Also, the use of configurations, fueling strategies, ignition and hydrocarbon fuels also extends the applicability of flameholding characteristics. These efforts will technology development to hypersonic cruise provide risk reduction for future flight demonstrator missions. In addition to the hydrocarbon-fueled engine development, specifically the RBCC and combustor development in the ground and flight TBCC engine concepts and will generate a demonstration projects previously discussed, there parametric design database for multi-speed is a need for fundamental supersonic combustion combustor optimization. These efforts are coupled studies for hydrogen-fueled engines, including an with investigations of combustion flow physics, 7 AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference AIAA2002-5188 using the CARS technique. 28The experiment was adopted as a test case by the RTO working group on scramjet propulsion (Working Group 10).29Near- term objectives of this work include further applications of optical diagnostic techniques to obtain simultaneous temperature and species measurements. The VULCAN CFD code, developed at NASA LaRC, is a Navier-Stokes code used to simulate chemically-reacting flow fields in scramjet combustors and flowpaths. 3f Previously, upgrades to this code, including volume grid capabilities, physical models and convergence acceleration i.....i., Vitiated Air Stream Combustor Model enhancements were incorporated. The combustor data sets described here comprise a partial data base for CFD code validation for this class of flows. Additional work has included the use of ......_.. Facility Nozzle /"-_"-"_ laser-based diagnostics along with flow seeding to ' Divergent Section obtain fuel plume images and velocity measurements. Reference 31 described the use of Figure 8. Langley Direct-Connect Supersonic this technique to obtain measurements in a Mach 2 Combustion Test Facility (DCSCTF). hydrogen-air combustor with a 10° unswept ramp computational modeling and associated diagnostic fuel injector. The VULCAN CFD code was used to tool development (described in the next section). obtain flow field predictions of this flowpath and Far-term goals are to design and demonstrate a comparisons with experimental wall pressures and flight-weight, regeneratively-cooled, hydrogen-fueled fuel plume images are shown. These comparisons combustor for future Mach 0-15 flight demonstration indicate an underprediction in the level of turbulent and vision vehicle development. mixing and heat release due to combustion. This work suggests that improvements in turbulence Analysis Tool and Diaqnostic Development modeling and turbulent-chemistry interactions are Computational Fluid Dynamics (CFD) needed to improve supersonic combustor modeling modeling of scramjet combustors is complicated by inthis speed range. various physical processes, including large regions of subsonic flow, separated flow regions, complex Future efforts are expected to focus on the mixing phenomena, non-equilibrium transfer of development of validated nose-to-tail design and turbulence energy, and interactions between analysis tools. This includes advanced physical turbulence and chemical kinetics that may impact modeling capabilities for turbulence, turbulence- both the chemical reactions and turbulence field. 26 chemistry interactions and reduced combustion Limitations in physical modeling capabilities as well kinetics models for hydrocarbon and hydrogen as computational overhead costs limit the practical combustion. Ultimately, diagnostic tool development use of three-dimensional CFD tools in design and will enable routine measurements of all combustor development of scramjet combustors. Several efforts _ MEA,_L_RE_,_ ENT L,DqAT_QN_ are underway to incorporate advanced physical I I models and algorithmic enhancements in CFD tools I I and to acquire high-fidelity data sets for code calibration. The Coherent anti-Stokes Raman Spectroscopy (CARS) technique has been used recently to acquire flow-field data in the LaRC DCSCTF for this purpose. 28'29These efforts have succeeded in mapping mean and RMS temperature fluctuations in a supersonic combustor. Figure 9 Figure 9. Supersonic Combustor (SCHOLAR) shows a generic supersonic combustor model with Model. angular fuel injection used for these studies and figure 10 shows mean temperature maps developed flow field parameters (temperature, velocity, species 8 AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference AIAA 2002-5188 engine control mechanisms. Engine control algorithms are required to control the engine variable contraction ratio schedule and control fuel flow rates to provide required thrust to meet mission objectives, enable mode transitions and to prevent and recover from engine unstart and flameout. These issues were studied in the Hyper-X program for the Mach 7 and 10 flight experiments and iL...ii,iii_i__ contributed to the development of flight control laws for the X-43A vehicles. This research is being further extended to the Mach 4-7 range and matured for future flight demonstrator vehicle development. iiiiiiiiiiiiiiiii_i_ _ililililii_licl__i_i_ Near-term project plans consist of additional freejet testing efforts in the Langley CHSTF and the AHSTF. First, the existing Mach 5 dual-fuel experimental engine (DFX) will be tested in the CHSTF to extend the low-speed (Mach 3-5) hydrogen-fueled engine performance database for DMSJ engines and to further examine facility test- Figure 10. Mean temperature maps obtained in gas vitiation effects. Second, further testing of the the DCSCTF using CARS (ref. 25). HXEM will be conducted in the AHSTF to investigate control laws for scramjet engine operation over the concentrations) to provide the necessary databases Mach 4-7 range. The HXEM testing will take place to fully characterize combustor flow fields and to following the installation of a new, high-quality flow, validate CFD methods. Mach 6 facility nozzle in the tunnel circuit. A thorough calibration at three total enthalpy levels will DMSJ/Combined-Cycle Performance and Operability be conducted followed by HXEM performance Dual-mode scramjet engine cycles must testing to examine the effects of nozzle exit flow function over a wide Mach number range in order to uniformity on engine performance. meet mission requirements. Current ground test databases exist only over a limited Mach number A longer-term objective of the program is to range. For candidate X-43B demonstrator conduct a comprehensive test program of a variable- configurations, engine performance, operability and geometry dual-mode scramjet engine configuration system weights must enable accomplishment of the over the Mach 3-8 speed range. The goal of this test Mach 7 mission objective. Future Mach 15 and program will be to investigate parametric SSTO operational concepts require scramjet performance at fixed contraction ratios, examine operation up to Mach 15 before transitioning to a contraction ratio changes during runs, examine rocket cycle or scramjet ejector cycle for orbital performance during mode transitions and to verify insertion. Therefore, mission specific impulse values closed-loop engine control algorithms. Real-time must be maximized over the Mach number range in enthalpy and dynamic pressure variation during order to provide acceptable performance margin. tests will be accomplished. Variable-geometry demonstration with heat-sink hardware will be Historically, a number of engine test followed with flight-like, regeneratively-cooled, programs at NASA-Langley, discussed previously, hydrogen-fueled engine ground testing. have contributed to the performance database for airframe-integrated scramjet flowpaths. 3'4 Hypervelocity Scram jet Development Depending on the Mach number range and specific As indicated in figure 1, the specific impulse mission requirements, efficient inlet operation over of the scramjet cycle decreases as Mach number the applicable flight regime may necessitate increases. Heat release due to combustion is variable contraction ratios. Variable-geometry inversely proportional to the square of the concepts have been examined to provide this freestream Mach number. At Mach 15, the required operability. combustion energy is approximately less than 25- percent of the free stream kinetic energy, accounting Efficient multi-speed engine operation also for flow field losses. 9At these Mach numbers, small requires the development of fueling strategies and changes in effective specific impulse can cause 9 AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference AIAA2002-5188 significant changes in vehicle take-off gross weights, to assess facility test conditions. thus impacting the ability of the system to meet mission performance requirements. 32 This Near-term efforts also consist of represents a practical upper limit for efficient development of a comprehensive Mach 12-15 scramjet engine operation without LOX- performance database. This will enable the design augmentation. A further understanding of the of a future flight demonstrator to validate Mach 15 fundamental physical processes that govern engine scramjet performance. Long-term goals include performance in the hypervelocity speed range is demonstrations of a flight-weight combustor at Mach required in order to optimize flowpath lines for 15 conditions and LOX-augmented ejector-scramjet efficient operation. cycles to enable orbital insertion for SSTO systems. System studies indicate that LOX-augmentation may At hypervelocity speeds, scramjet flow be required for efficient orbital insertion in SSTO physics are characterized by very short residence airbreathing launch vehicles. 37 times. Robust flowpath design is dependent on an understanding of the following flow-field and Propulsion-Airframe Inteqration (PAl) combustion phenomena: fuel injector geometry, Hypersonic airbreathing vehicles are mixing, flameholding, and combustion efficiency as characterized by highly integrated systems with a well as thermal balance and protection high degree of interaction between the airframe and requirements. The NASA-Langley HYPULSE facility, propulsion flowpath. Aerothermodynamic located at and operated by Allied Aerospace, Inc. performance cannot be decoupled from propulsion (GASL Division) has been used to perform performance, due to shared surfaces and flow field hypervelocity scramjet research in a shock- interactions, as depicted in figure 12. Therefore, a expansion-tunnel (SET) mode at enthalpy levels significant challenge in the design and development duplicating Mach numbers above 10. 33-35 A of hypersonic airbreathing flight vehicles is the schematic of the NASA HYPULSE Facility is shown determination of aero-propulsive interactions and in figure 11. Tests have been conducted most installed vehicle performance. Design tools and recently on a scramjet flowpath model, ground testing techniques are needed to fully representative of the Mach 10 X-43A scramjet characterize these effects across the applicable flowpath lines, at Mach 15 conditions. 36 speed ranges. Future test technique development will focus During the NASP program, significant work on a definition and calibration of HYPULSE (SET) at was done to investigate the use of cold-gas mixtures baseline test points in the Mach 12-15 range. This to simulate powered scramjet exhaust products in 38 40 includes the design, using CFD, and fabrication of a ground test facilities. - This technique was facility nozzle suitable for scramjet engine tests, and investigated in the supersonic and hypersonic speed efforts to optimize and calibrate the shock tunnel exit regimes (Mach 4-10) with powered metric aftbody flow conditions for these flight Mach numbers. models to develop the technique and measure Various flow diagnostic techniques, including external nozzle pressures, exhaust plume schlieren, fuel-plume (planar) imaging, water impingement on wing surfaces and aftbody forces temperature and concentration by laser absorption and moments. Analysis to examine the correlation of and laser holographic interferometry, will be applied cold simulant gases to hot combustion products was initiated, but not completed due to the termination of this program. The X-43A flight project undertook a significant ground testing and computational effort to 41 42 build the pre-flight vehicle database. ' This effort consisted of un-powered aerothermodynamic testing with powered force and moment increments supplied by CFD predictions. These predictions were verified by full-flowpath force and moment data obtained from the HXFE testing in the Langley 8-Ft. HTT. 18 Future NASA efforts in this area will build on Figure 11. NASA HYPULSE Facility. NASP and X-43A research and will focus on test ]0 AIAA/AAAF International Space Planes and Hypersonic Systems and Technologies Conference

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