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Sample records for airframe-integrated scramjet engine

  1. Fabrication and development of several heat pipe honeycomb sandwich panel concepts. [airframe integrated scramjet engine

    NASA Technical Reports Server (NTRS)

    Tanzer, H. J.

    1982-01-01

    The feasibility of fabricating and processing liquid metal heat pipes in a low mass honeycomb sandwich panel configuration for application on the NASA Langley airframe-integrated Scramjet engine was investigated. A variety of honeycomb panel facesheet and core-ribbon wick concepts was evaluated within constraints dictated by existing manufacturing technology and equipment. The chosen design consists of an all-stainless steel structure, sintered screen facesheets, and two types of core-ribbon; a diffusion bonded wire mesh and a foil-screen composite. Cleaning, fluid charging, processing, and process port sealing techniques were established. The liquid metals potassium, sodium and cesium were used as working fluids. Eleven honeycomb panels 15.24 cm X 15.24 cm X 2.94 cm were delivered to NASA Langley for extensive performance testing and evaluation; nine panels were processed as heat pipes, and two panels were left unprocessed.

  2. SCRAM - AN ENGINEER'S TOOL FOR PREDICTION OF AIRFRAME INTEGRATED SCRAMJET PERFORMANCE

    NASA Technical Reports Server (NTRS)

    Walton, J. T.

    1994-01-01

    This program determines the one-dimensional performance for an airframe integrated supersonic combustion ramjet (scramjet). The supersonic combustion ramjet cycle, which uses hydrogen for fuel and atmospheric air for oxidation, is essential for the development of a propulsion system for single-stage-to-orbit aerospace vehicles. These vehicles are intended to be launched horizontally, as opposed to vertical launching for current space vehicles. In addition, they must achieve hypersonic flight to Mach 25 prior to orbital insertion into low Earth orbit. The propulsion system of these vehicles must be reusable, efficient, and cost effective. The scramjet cycle analysis code performs nose-to-tail, hydrogen fueled, Airframe Integrated Scramjet (AIS) simulation in a real gas flow with equilibrium thermodynamic properties. This allows ready generation of preliminary estimates for SCRAM cycle performance. SCRAM is a reliable, efficient, and speedy design tool that is useable on all standard computers down to IBM PC-AT compatible machines. Developed in the Hypersonic Propulsion Branch at NASA Langley Research Center for the Hypersonic Research Engine and Langley 3-Strut engine programs, the current version of this code has been modified by the NASA Dryden Flight Research Facility of the Ames Research Center for the purpose of supporting the Langley Strutless Parametric engine and National AeroSpace Plane (NASP) engine test programs. The current version of SCRAM optimizes the tradeoffs between the needs for computational speed, accuracy, and future modifications. The program utilizes a five station geometry model, with variable step size between each station, to analyze a vehicle nose-to-tail mass capture stream tube control-volume with real gas equilibrium flow properties. SCRAM applies the laws of Conservation of Mass, Momentum, and Energy across each step to calculate the changing flow parameters along the control volume. The code incorporates an integral boundary layer

  3. Subscale, hydrogen-burning, airframe-integrated-scramjet: Experimental and theoretical evaluation of a water cooled strut airframe-integrated-scramjet: Experimental leading edge

    NASA Technical Reports Server (NTRS)

    Pinckney, S. Z.; Guy, R. W.; Beach, H. L., Jr.; Rogers, R. C.

    1975-01-01

    A water-cooled leading-edge design for an engine/airframe integrated scramjet model strut leading edge was evaluated. The cooling design employs a copper cooling tube brazed just downstream of the leading edge of a wedge-shaped strut which is constructed of oxygen-free copper. The survival of the strut leading edge during a series of tests at stagnation point heating rates confirms the practicality of the cooling design. A finite difference thermal model of the strut was also proven valid by the reasonable agreement of calculated and measured values of surface temperature and cooling-water heat transfer.

  4. Thermal design and analysis of a hydrogen-burning wind tunnel model of an airframe-integrated scramjet

    NASA Technical Reports Server (NTRS)

    Guy, R. W.; Mueller, J. N.; Pinckney, S. Z.; Lee, L. P.

    1976-01-01

    An aerodynamic model of a hydrogen burning, airframe integrated scramjet engine has been designed, fabricated, and instrumented. This model is to be tested in an electric arc heated wind tunnel at an altitude of 35.39 km (116,094 ft.) but with an inlet Mach number of 6 simulating precompression on an aircraft undersurface. The scramjet model is constructed from oxygen free, high conductivity copper and is a heat sink design except for water cooling in some critical locations. The model is instrumented for pressure, surface temperature, heat transfer rate, and thrust measurements. Calculated flow properties, heat transfer rates, and surface temperature distributions along the various engine components are included for the conditions stated above. For some components, estimates of thermal strain are presented which indicate significant reductions in plastic strain by selective cooling of the model. These results show that the 100 thermal cycle life of the engine was met with minimum distortion while staying within the 2669 N (600 lbf) engine weight limitation and while cooling the engine only in critical locations.

  5. Engine-Airframe Integration for Rotorcraft.

    DTIC Science & Technology

    1986-05-01

    maximum efficien- cy in terms of the range of inlet to free stream velocity ratio (Vi/Vo) as 0.0 Vi/Vo 6 0.65, within which local velocity gradients...ft r = radius, ft Subscripts a = air i = inner surface 1-9 o = outer surface c = cowl e = engine - = outside or free stream air 1 = inlet stations 2...these limift of L/D , the designer has the option of nozzling the primary exit area to induce a higher pAmary stream momentum, or adding a diffuser

  6. Internal Shock Interactions in Propulsion/Airframe Integrated Three-Dimensional Sidewall Compression Scramjet Inlets

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.; Perkins, John N.

    1992-01-01

    The advantages and design requirements of propulsion/airframe integration for high Mach number flight have led to extensive study of the three-dimensional sidewall compression scramjet inlet in recent years. Recent research publications have indicated testing over a broad range of Mach number (2 to 18) in a variety of test gases, such as air, helium, and tetrafluoromethane. Multiple experimental techniques have been employed to obtain detailed internal shock interaction data, performance data, and inlet starting limits. Computational fluid dynamics has been effectively used for preliminary parametric studies as well as in parallel with experiments to aid in the explanation of unusual or unexpected flow phenomena. Inlets of this genre afford a relatively simple, generic geometry while producing a highly complex, three-dimensional flow field dominated by shock/shock and shock/boundary layer interactions. While the importance of the viscous effects in high speed inlet interactions is recognized, the present work addresses in a parametric fashion the inviscid effects of leading edge sweep, sidewall compression, and inflow Mach number on the internal shock structure in terms of inlet compression and mass capture. In the process, the source of the of the Mach number invariance with leading edge sweep for a constant sidewall compression class of inlet is identified, and a previously undocumented spillage phenomenon in a constant effective wedge angle class of inlets is discussed.

  7. Advanced general aviation comparative engine/airframe integration study

    NASA Technical Reports Server (NTRS)

    Huggins, G. L.; Ellis, D. R.

    1981-01-01

    The NASA Advanced Aviation Comparative Engine/Airframe Integration Study was initiated to help determine which of four promising concepts for new general aviation engines for the 1990's should be considered for further research funding. The engine concepts included rotary, diesel, spark ignition, and turboprop powerplants; a conventional state-of-the-art piston engine was used as a baseline for the comparison. Computer simulations of the performance of single and twin engine pressurized aircraft designs were used to determine how the various characteristics of each engine interacted in the design process. Comparisons were made of how each engine performed relative to the others when integrated into an airframe and required to fly a transportation mission.

  8. Thermal-structural Design Study of an Airframe-integrated Scramjet

    NASA Technical Reports Server (NTRS)

    Buchmann, O. A.

    1979-01-01

    Design concepts for the cooled structures assembly for the Langley Scramjet engine, for engine subsystems, and for the aircraft/engine interface were developed and evaluated. Results show that the objectives for the Scramjet engine can be met. A thermal protection system was defined that makes it possible to attain a life of 100 hours and 1000 cycles, which is the specified goal. With stoichiometric combustion, the fuel provides an adequate heat sink for cooling the engine at Mach numbers up to 9 at the minimum fuel flow condition. The mechanical design is feasible for manufacture using conventional materials. For the cooled structures in a six-module engine, the mass per unit capture area is 1328 kg/sq m (259 lb/sq ft). The total mass of a six-module engine assembly including the fuel system is 1577 kg (3477 lb).

  9. Thermal-structural design study of an airframe-integrated Scramjet

    NASA Technical Reports Server (NTRS)

    Killackey, J. J.; Katinsky, E. A.; Tepper, S.; Vuigner, A. A.

    1978-01-01

    Design concepts are developed and evaluated for a cooled structures assembly for the Scramjet engine, for engine subsystems mass, volume, and operating requirements, and for the aircraft/engine interface. A thermal protection system was defined that makes it possible to attain a life of 100 hours and 1000 cycles. The coolant equivalence ratio at the Mach 10 maximum thermal loading condition is 0.6, indicating a capacity for airframe cooling. The mechanical design is feasible for manufacture using conventional materials. For the cooled structures in a six-module engine, the mass per unit capture area is 12.4 KN/sq m. The total weight of a six-module engine assembly including the fuel system is 14.73 KN.

  10. The Engineering Design of Engine/Airframe Integration for the SAENGER Fully Reusable Space Transportation System

    DTIC Science & Technology

    2010-09-01

    propellant rocket engine 26. Parachute container 31. RAMJET nitrogen cooling system storage tank 32. RAMJET control unit 33. High- pressure helium bottle...33. High- pressure helium bottle 34. RAMJET and its service lines fairing 35. Thermal-insulated LH2 tank 36. Studied RAMJET module 37. RAMJE fuel...operational modes and the results of the engine/airframe integration. Major experimental (EFD) and computational ( CFD ) results from the detailed intake

  11. Advanced general aviation engine/airframe integration study

    NASA Technical Reports Server (NTRS)

    Zmroczek, L. A.

    1982-01-01

    A comparison of the in-airframe performance and efficiency of the advanced engine concepts is presented. The results indicate that the proposed advanced engines can significantly improve the performance and economy of general aviation airplanes. The engine found to be most promising is the highly advanced version of a rotary combustion (Wankel) engine. The low weight and fuel consumption of this engine, as well as its small size, make it suited for aircraft use.

  12. Propulsion System Airframe Integration Issues and Aerodynamic Database Development for the Hyper-X Flight Research Vehicle

    NASA Technical Reports Server (NTRS)

    Engelund, Walter C.; Holland, Scott D.; Cockrell, Charles E., Jr.; Bittner, Robert D.

    1999-01-01

    NASA's Hyper-X Research Vehicle will provide a unique opportunity to obtain data on an operational airframe integrated scramjet propulsion system at true flight conditions. The airframe integrated nature of the scramjet engine with the Hyper-X vehicle results in a strong coupling effect between the propulsion system operation and the airframe s basic aerodynamic characteristics. Comments on general airframe integrated scramjet propulsion system effects on vehicle aerodynamic performance, stability, and control are provided, followed by examples specific to the Hyper-X research vehicle. An overview is provided of the current activities associated with the development of the Hyper-X aerodynamic database, including wind tunnel test activities and parallel CFD analysis efforts. A brief summary of the Hyper-X aerodynamic characteristics is provided, including the direct and indirect effects of the airframe integrated scramjet propulsion system operation on the basic airframe stability and control characteristics.

  13. Study of hypersonic propulsion/airframe integration technology

    NASA Technical Reports Server (NTRS)

    Hartill, W. R.; Goebel, T. P.; Vancamp, V. V.

    1978-01-01

    An assessment is done of current and potential ground facilities, and analysis and flight test techniques for establishing a hypersonic propulsion/airframe integration technology base. A mach 6 cruise prototype aircraft incorporating integrated Scramjet engines was considered the baseline configuration, and the assessment focused on the aerodynamic and configuration aspects of the integration technology. The study describes the key technology milestones that must be met to permit a decision on development of a prototype vehicle, and defines risk levels for these milestones. Capabilities and limitations of analysis techniques, current and potential ground test facilities, and flight test techniques are described in terms of the milestones and risk levels.

  14. Novel inlet-airframe integration methodology for hypersonic waverider vehicles

    NASA Astrophysics Data System (ADS)

    Ding, Feng; Liu, Jun; Shen, Chi-bing; Huang, Wei

    2015-06-01

    With the aim of integrating a ramjet or scramjet with an airframe, a novel inlet-airframe integration methodology for the hypersonic waverider vehicle is proposed. For this newly proposed design concept and for the specified flight conditions, not only the forebody of the vehicle but also its engine cowl and wings can ride on the bow shock wave, causing the bow shock wave to remain attached to the leading edge for the entire length of the vehicle. Thus, this integrated vehicle can take full advantage of the waverider's high lift-to-drag ratio characteristics and the ideal pre-compression surface for the engine. In this work, a novel inlet-airframe integrated axisymmetric basic flow model that accounts for both external and internal flows is first designed using the method of characteristics and the streamline tracing technique. Subsequently, the design of the inlet-airframe integrated waverider vehicle is generated from the integrated axisymmetric basic flow model using the streamline tracing technique. Then, the design methodologies of both the integrated axisymmetric basic flow model and the integrated waverider vehicle are verified by a computational numerical method. Finally, the viscous effects and performance of both the integrated axisymmetric basic flow model and the integrated waverider vehicle are analysed under the design condition using the numerical computation. The obtained results show that the proposed approach is effective in designing the integrated hypersonic waverider vehicle.

  15. Engineering Engine/Airframe Integration for Fully Reusable Space Transportation Systems

    DTIC Science & Technology

    2010-09-01

    system concept for SÄNGER first stage is important to compare the major performance characteristics of the engine cycle types which were under...AUTHOR(S) 5d. PROJECT NUMBER 5e. TASK NUMBER 5f. WORK UNIT NUMBER 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) Aerospace Consulting 82024...Taufkirchen 8. PERFORMING ORGANIZATION REPORT NUMBER 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSOR/MONITOR’S ACRONYM(S) 11

  16. Small Engine Technology (SET). Task 33: Airframe, Integration, and Community Noise Study

    NASA Technical Reports Server (NTRS)

    Lieber, Lys S.; Elkins, Daniel; Golub, Robert A. (Technical Monitor)

    2002-01-01

    Task Order 33 had four primary objectives as follows: (1) Identify and prioritize the airframe noise reduction technologies needed to accomplish the NASA Pillar goals for business and regional aircraft. (2) Develop a model to estimate the effect of jet shear layer refraction and attenuation of internally generated source noise of a turbofan engine on the aircraft system noise. (3) Determine the effect on community noise of source noise changes of a generic turbofan engine operating from sea level to 15,000 feet. (4) Support lateral attenuation experiments conducted by NASA Langley at Wallops Island, VA, by coordinating opportunities for Contractor Aircraft to participate as a noise source during the noise measurements. Noise data and noise prediction tools, including airframe noise codes, from the NASA Advanced Subsonic Technology (AST) program were applied to assess the current status of noise reduction technologies relative to the NASA pillar goals for regional and small business jet aircraft. In addition, the noise prediction tools were applied to evaluate the effectiveness of airframe-related noise reduction concepts developed in the AST program on reducing the aircraft system noise. The AST noise data and acoustic prediction tools used in this study were furnished by NASA.

  17. MHD Energy Bypass Scramjet Engine

    NASA Technical Reports Server (NTRS)

    Mehta, Unmeel B.; Bogdanoff, David W.; Park, Chul; Arnold, Jim (Technical Monitor)

    2001-01-01

    Revolutionary rather than evolutionary changes in propulsion systems are most likely to decrease cost of space transportation and to provide a global range capability. Hypersonic air-breathing propulsion is a revolutionary propulsion system. The performance of scramjet engines can be improved by the AJAX energy management concept. A magneto-hydro-dynamics (MHD) generator controls the flow and extracts flow energy in the engine inlet and a MHD accelerator downstream of the combustor accelerates the nozzle flow. A progress report toward developing the MHD technology is presented herein. Recent theoretical efforts are reviewed and ongoing experimental efforts are discussed. The latter efforts also include an ongoing collaboration between NASA, the US Air Force Research Laboratory, US industry, and Russian scientific organizations. Two of the critical technologies, the ionization of the air and the MHD accelerator, are briefly discussed. Examples of limiting the combustor entrance Mach number to a low supersonic value with a MHD energy bypass scheme are presented, demonstrating an improvement in scramjet performance. The results for a simplified design of an aerospace plane show that the specific impulse of the MHD-bypass system is better than the non-MHD system and typical rocket over a narrow region of flight speeds and design parameters. Equilibrium ionization and non-equilibrium ionization are discussed. The thermodynamic condition of air at the entrance of the engine inlet determines the method of ionization. The required external power for non-equilibrium ionization is computed. There have been many experiments in which electrical power generation has successfully been achieved by magneto-hydrodynamic (MHD) means. However, relatively few experiments have been made to date for the reverse case of achieving gas acceleration by the MHD means. An experiment in a shock tunnel is described in which MHD acceleration is investigated experimentally. MHD has several

  18. Hyper-X Research Vehicle - Artist Concept in Flight with Scramjet Engine Firing

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This is an artist's depiction of a Hyper-X research vehicle under scramjet power in free-flight following separation from its booster rocket. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need

  19. Heat pipe cooling for scramjet engines

    NASA Technical Reports Server (NTRS)

    Silverstein, Calvin C.

    1986-01-01

    Liquid metal heat pipe cooling systems have been investigated for the combustor liner and engine inlet leading edges of scramjet engines for a missile application. The combustor liner is cooled by a lithium-TZM molybdenum annular heat pipe, which incorporates a separate lithium reservoir. Heat is initially absorbed by the sensible thermal capacity of the heat pipe and liner, and subsequently by the vaporization and discharge of lithium to the atmosphere. The combustor liner temperature is maintained at 3400 F or less during steady-state cruise. The engine inlet leading edge is fabricated as a sodium-superalloy heat pipe. Cooling is accomplished by radiation of heat from the aft surface of the leading edge to the atmosphere. The leading edge temperature is limited to 1700 F or less. It is concluded that heat pipe cooling is a viable method for limiting scramjet combustor liner and engine inlet temperatures to levels at which structural integrity is greatly enhanced.

  20. Aerothermodynamic flow phenomena of the airframe-integrated supersonic combustion ramjet

    NASA Technical Reports Server (NTRS)

    Walton, James T.

    1992-01-01

    The unique component flow phenomena is discussed of the airframe-integrated supersonic combustion ramjet (scramjet) in a format geared towards new players in the arena of hypersonic propulsion. After giving an overview of the scramjet aerothermodynamic cycle, the characteristics are then covered individually of the vehicle forebody, inlet, combustor, and vehicle afterbody/nozzle. Attention is given to phenomena such as inlet speeding, inlet starting, inlet spillage, fuel injection, thermal choking, and combustor-inlet interaction.

  1. Combustion Characteristics of a Hypermixer Scramjet Engine

    NASA Astrophysics Data System (ADS)

    Sunami, Tetsuji; Itoh, Katsuhiro; Komuro, Tomoyuki; Sato, Kazuo

    A scramjet engine with a wall-mounted hypermixer injector, which generates streamwise vortices for enhancing supersonic mixing and combustion, is examined at a Mach 8 simulated flight condition in the High Enthalpy Shock Tunnel (HIEST). The engine and the fuel injector are full scale models of the HyShot-IV flight experiment planned for 2005 by JAXA and University of Queensland (UQ). Main purpose of the present study is to clarify the combustion and operation characteristics of the hypermixer scramjet owing to the ability of the streamwise vortices for mixing enhancement and boundary layer control. For comparison, two injectors with normal and parallel injection without streamwise vortex generation are also examined. The results show the superior performance of the hypermixer injector in scramjet mode obtaining higher pressure rise in a shorter distance compared to the other two injectors. In the case of the hypermixer injector, a 1D analysis of an inviscid nozzle flow shows the increment in the specific impulse due to combustion to be 2,649 and 2,224 sec for the equivalence ratio, Φ=0.3 and 0.6, respectively. At Φ=1.0 and 1.5, sudden rapid combustion of the premixed fuel at the end of the combustor generates a strong pressure wave, which propagates upstream up to the injector location and decades there. As a result, a new quasi-steady combusting flow is established throughout the combustor downstream of the injector. The pressure wave is identified as a kind of detonation wave, which is suggested to propagate upstream mainly through the streamwise vortices. As a driving force of the upstream propagation of the detonation wave, mixing and combustion enhanced through the interaction between the detonation wave and the streamwise vortices are considered.

  2. Internal performance predictions for Langley scramjet engine module

    NASA Technical Reports Server (NTRS)

    Pinckney, S. Z.

    1978-01-01

    A one dimensional theoretical method for the prediction of the internal performance of a scramjet engine is presented. The effects of changes in vehicle forebody flow parameters and characteristics on predicted thrust for the scramjet engine were evaluated using this method, and results are presented. A theoretical evaluation of the effects of changes in the scramjet engine's internal parameters is also presented. Theoretical internal performance predictions, in terms thrust coefficient and specific impulse, are provided for the scramjet engine for free stream Mach numbers of 5, 6, and 7 free stream dynamic pressure of 23,940 N/sq m forebody surface angles of 4.6 deg to 14.6 deg, and fuel equivalence ratio of 1.0.

  3. Hypersonic airbreathing propulsion/airframe integration

    NASA Technical Reports Server (NTRS)

    Weidner, John P.

    1992-01-01

    Recent interest in airbreathing hypersonic flight has centered around the need to develop advanced space launch systems which can reduce the cost of inserting payloads in orbit and make space more accessible. An effect of the thermal environment is to require the vehicle to operate at high altitudes, in very thin air, to maintain aircraft structural load limits. The high altitudes at which the hypersonic vehicle must operate give rise to the concept of an airframe integrated propulsion system to provide a much larger inlet and nozzle to process the required volume of air at low density, atmospheric conditions. In the integrated system, the forward portion of the vehicle compresses the air flow and serves as the external portion of the inlet; the aftbody completes the expansion process for the nozzle. In addition, the engine, which is contained between the body and the forebody shock wave, lends itself to a modular integration of a number of separate engines. In this manner, a relatively small engine can be defined to allow engine development in existing ground facilities.

  4. Hypersonic propulsion. [supersonic combustion ramjet engines

    NASA Technical Reports Server (NTRS)

    Beach, H. L., Jr.

    1979-01-01

    Research on hydrogen fueled scramjet engines for hypersonic flight is reviewed. Component developments, computational methods, and preliminary ground tests of subscale scramjet engine modules at Mach 4 and 7 are emphasized. Airframe integration, structures, and flow diagnostics are also discussed. It is shown that mixed-mode perpendicular and parallel fuel injection controls heat release over a wide Mach range and the fixed geometry inlet gives good performance over a wide range of Mach numbers.

  5. Airframe-integrated propulsion system for hypersonic cruise vehicles

    NASA Technical Reports Server (NTRS)

    Jones, R. A.; Huber, P. W.

    1978-01-01

    Research on a new, hydrogen burning, airbreathing engine concept which offers good potential for efficient hypersonic cruise vehicles is considered. Features of the engine which lead to good performance include; extensive engine-airframe integration, fixed geometry, low cooling, and the control of heat release in the supersonic combustor by mixed-modes of fuel injection from the combustor entrance. The engine concept is described along with results from inlet tests, direct-connect combustor tests, and tests of two subscale boiler-plate research engines presently underway at conditions which simulate flight at Mach 4 and 7.

  6. RAMSCRAM: A flexible ramjet/scramjet engine simulation program

    NASA Technical Reports Server (NTRS)

    Burkardt, Leo A.; Franciscus, Leo C.

    1990-01-01

    With the resurgence of interest in high supersonic and hypersonic flight there is a need to simulate airbreathing engines which may be used in this flight regime. To meet this requirement the RAMSCRAM code was developed. The code calculates 1-D flow properties at each component interface and the overall performance of the engine. It uses equilibrium thermodynamics which accounts for dissociation and allows for any fuel or combination of fuels. The program can simulate ramjet, scramjet, rocket, and ducted rocket engines.

  7. Heat pipe cooling for scramjet engines. Final report

    SciTech Connect

    Silverstein, C.C.

    1986-12-01

    Liquid metal heat pipe cooling systems have been investigated for the combustor liner and engine inlet leading edges of scramjet engines for a missile application. The combustor liner is cooled by a lithium-TZM molybdenum annular heat pipe, which incorporates a separate lithium reservoir. Heat is initially absorbed by the sensible thermal capacity of the heat pipe and liner, and subsequently by the vaporization and discharge of lithium to the atmosphere. The combustor liner temperature is maintained at 3400 F or less during steady-state cruise. The engine inlet leading edge is fabricated as a sodium-superalloy heat pipe. Cooling is accomplished by radiation of heat from the aft surface of the leading edge to the atmosphere. The leading edge temperature is limited to 1700 F or less. It is concluded that heat pipe cooling is a viable method for limiting scramjet combustor liner and engine inlet temperatures to levels at which structural integrity is greatly enhanced.

  8. Validation of scramjet exhaust simulation technique

    NASA Technical Reports Server (NTRS)

    Hopkins, H. B.; Konopka, W.; Leng, J.

    1976-01-01

    Scramjet/airframe integration design philosophy for hypersonic aircraft results in configurations having lower aft surfaces that serve as exhaust nozzles. There is a strong coupling between the exhaust plume and the aerodynamics of the vehicle, making accurate simulation of the engine exhaust mandatory. The experimental verification of the simulation procedure is described. The detonation tube simulator was used to produce an exact simulation of the scramjet exhaust for a Mach 8 flight condition. The pressure distributions produced by the exact exhaust flow were then duplicated by a cool mixture Argon and Freon 13B1. Such a substitute gas mixture validated by the detonation tube technique could be used in conventional wind tunnel tests. The results presented show the substitute gas simulation technique to be valid for shockless expansions.

  9. Active cooling design for scramjet engines using optimization methods

    NASA Technical Reports Server (NTRS)

    Scotti, Stephen J.; Martin, Carl J.; Lucas, Stephen H.

    1988-01-01

    A methodology for using optimization in designing metallic cooling jackets for scramjet engines is presented. The optimal design minimizes the required coolant flow rate subject to temperature, mechanical-stress, and thermal-fatigue-life constraints on the cooling-jacket panels, and Mach-number and pressure contraints on the coolant exiting the panel. The analytical basis for the methodology is presented, and results for the optimal design of panels are shown to demonstrate its utility.

  10. Active cooling design for scramjet engines using optimization methods

    NASA Technical Reports Server (NTRS)

    Scotti, Stephen J.; Martin, Carl J.; Lucas, Stephen H.

    1988-01-01

    A methodology for using optimization in designing metallic cooling jackets for scramjet engines is presented. The optimal design minimizes the required coolant flow rate subject to temperature, mechanical-stress, and thermal-fatigue-life constraints on the cooling-jacket panels, and Mach-number and pressure constraints on the coolant exiting the panel. The analytical basis for the methodology is presented, and results for the optimal design of panels are shown to demonstrate its utility.

  11. Combustion Enhancement in Scramjet-Operation of a RBCC Engine

    NASA Astrophysics Data System (ADS)

    Sadatake Tomioka, By; Ryohei Kobayashi; Murakami, Atsuo; Shuichi Ueda; Komuro, Tomoyuki; Katsuhiro Itoh, And

    Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket Based Combined Cycle engine) are expected to be the most effective propulsion system for Booster stage of space launch vehicles. At hypersonic regime, it will be operated at rather high rocket engine output for final acceleration with some Isp gains due to air-breathing effects. In this regime, attaining thrust at this high-speed regime becomes very difficult, so that parallel injection of the fuel for scramjet combustion is favorable as the momentum of the injection can contribute to the thrust production. Thus, embedded rocket chamber was supposed to the operated as fuel rich gas generator at very high output. This configuration was tested at simulated flight Mach number of 7-11 at High Enthalpy Shock Tunnel (HIEST) with detonation tube as the source of the simulated rocket exhaust. However, combustion of the residual fuel in the rocket exhaust with airflow could not be attained. Direct-connect combustor tests were performed to evaluate effectiveness of a combustion enhancement technique termed auxiliary injection, i.e., a portion of fuel to be directly injected into airflow to provide ignition source for the residual fuel. Results of both the engine model tests at HIEST and the direct-connect tests are summarized and presented, and modification to the engine model for combustion enhancement was proposed.

  12. Hyper-X Engine Testing in the NASA Langley 8-Foot High Temperature Tunnel

    NASA Technical Reports Server (NTRS)

    Huebner, Lawrence D.; Rock, Kenneth E.; Witte, David W.; Ruf, Edward G.; Andrews, Earl H., Jr.

    2000-01-01

    Airframe-integrated scramjet engine tests have 8 completed at Mach 7 in the NASA Langley 8-Foot High Temperature Tunnel under the Hyper-X program. These tests provided critical engine data as well as design and database verification for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe- integrated scramjet flight data. The first model tested was the Hyper-X Engine Model (HXEM), and the second was the Hyper-X Flight Engine (HXFE). The HXEM, a partial-width, full-height engine that is mounted on an airframe structure to simulate the forebody features of the X-43, was tested to provide data linking flowpath development databases to the complete airframe-integrated three-dimensional flight configuration and to isolate effects of ground testing conditions and techniques. The HXFE, an exact geometric representation of the X-43 scramjet engine mounted on an airframe structure that duplicates the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle base, was tested to verify the complete design as it will be flight tested. This paper presents an overview of these two tests, their importance to the Hyper-X program, and the significance of their contribution to scramjet database development.

  13. Numerical investigation of combustion field of hypervelocity scramjet engine

    NASA Astrophysics Data System (ADS)

    Zhang, Shikong; Li, Jiang; Qin, Fei; Huang, Zhiwei; Xue, Rui

    2016-12-01

    A numerical study of the ground testing of a hydrogen-fueled scramjet engine was undertaken using the commercial computational-fluid-dynamics code CFD++. The simulated Mach number was 12. A 7-species, 9-reaction-step hydrogen-air chemistry kinetics system was adopted for the Reynolds-averaged Navier-Stokes simulation. The two-equation SST turbulence model, which takes into account the wall functions, was used to handle the turbulence-chemistry interactions. The results were validated by experimentally measuring the wall pressure distribution, and the values obtained proved to be in good agreement. The flow pattern at non-reaction/reaction is presented, as are the results of analyzing the supersonic premix/non-premix flame structure, the reaction heat release distribution in different modes, and the change in the equivalence ratio. In this study, we realize the working mode of a hypervelocity engine and provide some suggestions for the combustion organization of the engine as well as offer insight into the potential for exploiting the processes of combustion and flow.

  14. Diode Laser-Based Detection of Combustor Instabilities with Application to a Scramjet Engine

    DTIC Science & Technology

    2010-02-01

    combustion instability has focused primarily on subsonic turbulent combustors [7]. NOx emission regulations on gas turbines have driven the use of...Optical Access Laser Beam Location ermocouple eholder) Static Pressure Taps (Open circles)holders Combustor w, direct-connect scramjet flowpath. 4.8 s...AFRL-RZ-WP-TP-2010-2055 DIODE LASER-BASED DETECTION OF COMBUSTOR INSTABILITIES WITH APPLICATION TO A SCRAMJET ENGINE (POSTPRINT) Campbell

  15. Development and fabrication of structural components for a scramjet engine

    NASA Technical Reports Server (NTRS)

    Buchmann, O. A.

    1990-01-01

    A program broadly directed toward design and development of long-life (100 hours and 1,000 cycles with a goal of 1,000 hours and 10,000 cycles) hydrogen-cooled structures for application to scramjets is presented. Previous phases of the program resulted in an overall engine design and analytical and experimental characterization of selected candidate materials and concepts. The latter efforts indicated that the basic life goals for the program can be reached with available means. The main objective of this effort was an integrated, experimental evaluation of the results of the previous program phases. The fuel injection strut was selected for this purpose, including fabrication development and fabrication of a full-scale strut. Testing of the completed strut was to be performed in a NASA-Langley wind tunnel. In addition, conceptual designs were formulated for a heat transfer test unit and a flat panel structural test unit. Tooling and fabrication procedures required to fabricate the strut were developed, and fabrication and delivery to NASA of all strut components, including major subassemblies, were completed.

  16. Numerical study of the performance of a model scramjet engine

    NASA Astrophysics Data System (ADS)

    Alhumadi, Ayad

    A computational parametric investigation was conducted to study the effect of variations to several geometric parameters on the performance of a two-dimensional model scramjet engine (square cross section area for 3-D model). Geometric parameters included backstep location, height, and angle and fuel injector angle, diameter, and location. Two- and three-dimensional geometries have been studied, using a finite-volume computational fluid dynamics (CFD) code (FLUENT) with structured grids with sizes between 50,000 and 90,000 cells for the two-dimensional geometry and with structured hexahedral grid sizes between 650,000 and 949,725 cells for the three-dimensional geometry. Otherwise, identical values of program inputs were utilized for the two- and three-dimensional simulations. Performance parameters investigated were combustion efficiency, thrust, pressure losses, and the equivalence ratio for the hydrogen-air combustor. A set of values for independent variables was identified which resulted in maximum thrust, minimum pressure losses, and an equivalence ratio close to unity.

  17. Aeroacoustics of Propulsion Airframe Integration: Overview of NASA's Research

    NASA Technical Reports Server (NTRS)

    Thomas, Russell H.

    2003-01-01

    The integration of propulsion and airframe is fundamental to the design of an aircraft system. Many considerations influence the integration, such as structural, aerodynamic, and maintenance factors. In regard to the acoustics of an aircraft, the integration can have significant effects on the net radiated noise. Whether an engine is mounted above a wing or below can have a significant effect on noise that reaches communities below because of shielding or reflection of engine noise. This is an obvious example of the acoustic effects of propulsion airframe installation. Another example could be the effect of the pylon on the development of the exhaust plume and on the resulting jet noise. In addition, for effective system noise reduction the impact that installation has on noise reduction devices developed on isolated components must be understood. In the future, a focus on the aerodynamic and acoustic interaction effects of installation, propulsion airframe aeroacoustics, will become more important as noise reduction targets become more difficult to achieve. In addition to continued fundamental component reduction efforts, a system level approach that includes propulsion airframe aeroacoustics will be required in order to achieve the 20 dB of perceived noise reduction envisioned by the long-range NASA goals. This emphasis on the aeroacoustics of propulsion airframe integration is a new part of NASA s noise research. The following paper will review current efforts and highlight technical challenges and approaches.

  18. Scramjet Inlets

    DTIC Science & Technology

    2010-09-01

    integration et gestion thermique) 14. ABSTRACT The supersonic combustion ramjet, or scramjet, is the engine cycle most suitable for sustained...and the stream thrust of the flow exiting the engine. The flow enters the engine at ambient conditions and at the flight velocity, so calculation of

  19. Transpiring Cooling of a Scram-Jet Engine Combustion Chamber

    NASA Technical Reports Server (NTRS)

    Choi, Sang H.; Scotti, Stephen J.; Song, Kyo D.; Ries,Heidi

    1997-01-01

    The peak cold-wall heating rate generated in a combustion chamber of a scram-jet engine can exceed 2000 Btu/sq ft sec (approx. 2344 W/sq cm). Therefore, a very effective heat dissipation mechanism is required to sustain such a high heating load. This research focused on the transpiration cooling mechanism that appears to be a promising approach to remove a large amount of heat from the engine wall. The transpiration cooling mechanism has two aspects. First, initial computations suggest that there is a reduction, as much as 75%, in the heat flux incident on the combustion chamber wall due to the transpirant modifying the combustor boundary layer. Secondly, the heat reaching the combustor wall is removed from the structure in a very effective manner by the transpirant. It is the second of these two mechanisms that is investigated experimentally in the subject paper. A transpiration cooling experiment using a radiant heating method, that provided a heat flux as high as 200 Btu/sq ft sec ( approx. 234 W/sq cm) on the surface of a specimen, was performed. The experiment utilized an arc-lamp facility (60-kW radiant power output) to provide a uniform heat flux to a test specimen. For safety reasons, helium gas was used as the transpirant in the experiments. The specimens were 1.9-cm diameter sintered, powdered-stainless-steel tubes of various porosities and a 2.54cm square tube with perforated multi-layered walls. A 15-cm portion of each specimen was heated. The cooling effectivenes and efficiencies by transpiration for each specimen were obtained using the experimental results. During the testing, various test specimens displayed a choking phenomenon in which the transpirant flow was limited as the heat flux was increased. The paper includes a preliminary analysis of the transpiration cooling mechanism and a scaling conversion study that translates the results from helium tests into the case when a hydrogen medium is used.

  20. Alleviation of Facility/Engine Interactions in an Open-Jet Scramjet Test Facility

    NASA Technical Reports Server (NTRS)

    Albertson, Cindy W.; Emami, Saied

    2001-01-01

    Results of a series of shakedown tests to eliminate facility/engine interactions in an open-jet scramjet test facility are presented. The tests were conducted with the NASA DFX (Dual-Fuel eXperimental scramjet) engine in the NASA Langley Combustion Heated Scramjet Test Facility (CHSTF) in support of the Hyper-X program, The majority of the tests were conducted at a total enthalpy and pressure corresponding to Mach 5 flight at a dynamic pressure of 734 psf. The DFX is the largest engine ever tested in the CHSTF. Blockage, in terms of the projected engine area relative to the nozzle exit area, is 81% with the engine forebody leading edge aligned with the upper edge of the facility nozzle such that it ingests the nozzle boundary layer. The blockage increases to 95% with the engine forebody leading edge positioned 2 in. down in the core flow. Previous engines successfully tested in the CHSTF have had blockages of no more than 51%. Oil flow studies along with facility and engine pressure measurements were used to define flow behavior. These results guided modifications to existing aeroappliances and the design of new aeroappliances. These changes allowed fueled tests to be conducted without facility interaction effects in the data with the engine forebody leading edge positioned to ingest the facility nozzle boundary layer. Interaction effects were also reduced for tests with the engine forebody leading edge positioned 2 in. into the core flow, however some interaction effects were still evident in the engine data. A new shroud and diffuser have been designed with the goal of allowing fueled tests to be conducted with the engine forebody leading edge positioned in the core without facility interaction effects in the data. Evaluation tests of the new shroud and diffuser will be conducted once ongoing fueled engine tests have been completed.

  1. Surface cooling of scramjet engine inlets using heat pipe, transpiration, and film cooling

    SciTech Connect

    Modlin, J.M.; Colwell, G.T. Georgia Institute of Technology, Atlanta )

    1992-09-01

    This article reports the results of applying a finite-difference-based computational technique to the problem of predicting the transient thermal behavior of a scramjet engine inlet exposed to a typical hypersonic flight aerodynamic surface heating environment, including type IV shock interference heating. The leading-edge cooling model utilized incorporates liquid metal heat pipe cooling with surface transpiration and film cooling. Results include transient structural temperature distributions, aerodynamic heat inputs, and surface coolant distributions. It seems that these cooling techniques may be used to hold maximum skin temperatures to near acceptable values during the severe aerodynamic and type IV shock interference heating effects expected on the leading edge of a hypersonic aerospace vehicle scramjet engine. 15 refs.

  2. Upper-stage space shuttle propulsion by means of separate scramjet and rocket engines

    NASA Technical Reports Server (NTRS)

    Franciscus, L. C.; Allen, J. L.

    1972-01-01

    A preliminary mission study of a reusable vehicle from staging to orbit indicates payload advantages for a dual-propulsion system consisting of separate scramjet and rocket engines. In the analysis the scramjet operated continuously and the initiation of rocket operation was varied. For a stage weight of 500,000 lb the payload was 10.4 percent of stage weight or 70 percent greater than that of a comparable all-rocket-powered stage. When compared with a reusable two-state rocket vehicle having 50,000 lb payload, the use of the dual propulsion system for the second stage resulted in significant decreases in lift-off weight and empty weight, indicating possible lower hardware costs.

  3. Recent progress in scramjet/combined cycle engines at JAXA, Kakuda space center

    NASA Astrophysics Data System (ADS)

    Hiraiwa, Tetsuo; Ito, Katsuhiro; Sato, Shigeru; Ueda, Shuichi; Tani, Kouichiro; Tomioka, Sadatake; Kanda, Takeshi

    2008-09-01

    This report presents recent research activities of the Combined Propulsion Research Group of Japan Aerospace Exploration Agency. Aerodynamics and combustion of the scramjet engine and the rocket-ramjet combined-cycle engine, structure and material for the two engines and thermo-aerodynamic of a re-entry vehicle are major subjects. In Mach 6 condition tests, a scramjet engine model produced about 2000 N net thrust, whereas a model produced thrust almost equal to its drag in Mach 12 condition. A flight test of a combustor model was conducted with Hyshot-IV. A rocket-ramjet combined-cycle engine model is under construction with validation of the rocket engine component. Studies of combustor models and aerodynamic component models were conducted for demonstration of the engine operation and improvement of its performances. Light-weight cooling panel by electrochemical etching examined and C/ C composite structure was tested. Thermo-aerodynamics of re-entry vehicle was investigated and oxygen molecular density was measured also in high enthalpy flow.

  4. Results of the NASP Ames Integrated Mixing Hypersonic Engine (AIMHYE) Scramjet Test Program

    NASA Technical Reports Server (NTRS)

    Cavolowsky, John A.; Loomis, Mark P.; Deiwert, George S.

    1995-01-01

    This paper describes the test techniques and results from the National Aerospace Plane Government Work Package 53, the Ames Integrated Mixing Hypersonic Engine (AIMHYE) Scramjet Test program conducted in the NASA Ames 16-Inch Combustion Driven Shock Tunnel. This was a series of near full-scale scramjet combustor tests with the objective to obtain high speed combustor and nozzle data from an engine with injector configurations similar to the NASP E21 and E22a designs. The experimental test approach was to use a large combustor model (80-100% throat height) designed and fabricated for testing in the semi-free jet mode. The conditions tested were similar to the "blue book" conditions at Mach 12, 14, and 16. GWP 53 validated use of large, long test time impulse facilities, specifically the Ames 16-Inch Shock Tunnel, for high Mach number scramjet propulsion testing an integrated test rig (inlet, combustor, and nozzle). Discussion of key features of the test program will include: effects of the 2-D combustor inlet pressure profile; performance of large injectors' fueling system that included nozzlettes, base injection, and film cooling; and heat transfer measurements to the combustor. Significant instrumentation development and application efforts include the following: combustor force balance application for measurement of combustor drag for comparison with integrated point measurements of skin friction; nozzle metric strip for measuring thrust with comparison to integrated pressure measurements; and nonintrusive optical fiber-based diode laser absorption measurements of combustion products for determination of combustor performance. Direct measurements will be reported for specific test article configurations and compared with CFD solutions.

  5. Perspective of laser-induced plasma ignition of hydrocarbon fuel in Scramjet engine

    NASA Astrophysics Data System (ADS)

    Yang, Leichao; Li, Xiaohui; Liang, Jianhan; Yu, Xin; Li, Xipeng

    2016-01-01

    Laser-induced plasma ignition of an ethylene fuelled cavity was successfully conducted in a model scramjet engine combustor. The ethylene was injected 10mm upstream of cavity flameholder from 3 orifices 60 degree inclined relative to freestream direction. The 1064nm laser beam, from a Q-switched Nd:YAG laser source running at 3Hz and 200mJ per pulse, was focused into cavity for ignition. High speed photography was used to capture the transient ignition process. The laser-induced gas breakdown, flame kernel generation and propagation were all recorded and ensuing stable supersonic combustion was established in cavity. The flame kernel is found rotating anti-clockwise and gradually moves upwards as the entrainment of circulation flow in cavity. The flame is then stretched from leading edge to trailing edge to fully fill the entire cavity. Eventually, a stable combustion is achieved roughly 900μs after the laser pulse. The results show promising potentials for practical application. The perspective of laser-induced plasma ignition of hydrocarbon fuel in scramjet engine is outlined.

  6. Thermodynamic Analysis of Dual-Mode Scramjet Engine Operation and Performance

    NASA Technical Reports Server (NTRS)

    Riggins, David; Tacket, Regan; Taylor, Trent; Auslender, Aaron

    2006-01-01

    Recent analytical advances in understanding the performance continuum (the thermodynamic spectrum) for air-breathing engines based on fundamental second-law considerations have clarified scramjet and ramjet operation, performance, and characteristics. Second-law based analysis is extended specifically in this work to clarify and describe the performance characteristics for dual-mode scramjet operation in the mid-speed range of flight Mach 4 to 7. This is done by a fundamental investigation of the complex but predictable interplay between heat release and irreversibilities in such an engine; results demonstrate the flow and performance character of the dual mode regime and of dual mode transition behavior. Both analytical and computational (multi-dimensional CFD) studies of sample dual-mode flow-fields are performed in order to demonstrate the second-law capability and performance and operability issues. The impact of the dual-mode regime is found to be characterized by decreasing overall irreversibility with increasing heat release, within the operability limits of the system.

  7. Parametric scramjet analysis

    NASA Astrophysics Data System (ADS)

    Choi, Jongseong

    The performance of a hypersonic flight vehicle will depend on existing materials and fuels; this work presents the performance of the ideal scramjet engine for three different combustion chamber materials and three different candidate fuels. Engine performance is explored by parametric cycle analysis for the ideal scramjet as a function of material maximum service temperature and the lower heating value of jet engine fuels. The thermodynamic analysis is based on the Brayton cycle as similarly employed in describing the performance of the ramjet, turbojet, and fanjet ideal engines. The objective of this work is to explore material operating temperatures and fuel possibilities for the combustion chamber of a scramjet propulsion system to show how they relate to scramjet performance and the seven scramjet engine parameters: specific thrust, fuel-to-air ratio, thrust-specific fuel consumption, thermal efficiency, propulsive efficiency, overall efficiency, and thrust flux. The information presented in this work has not been done by others in the scientific literature. This work yields simple algebraic equations for scramjet performance which are similar to that of the ideal ramjet, ideal turbojet and ideal turbofan engines.

  8. Process-based Cost Estimation for Ramjet/Scramjet Engines

    NASA Technical Reports Server (NTRS)

    Singh, Brijendra; Torres, Felix; Nesman, Miles; Reynolds, John

    2003-01-01

    Process-based cost estimation plays a key role in effecting cultural change that integrates distributed science, technology and engineering teams to rapidly create innovative and affordable products. Working together, NASA Glenn Research Center and Boeing Canoga Park have developed a methodology of process-based cost estimation bridging the methodologies of high-level parametric models and detailed bottoms-up estimation. The NASA GRC/Boeing CP process-based cost model provides a probabilistic structure of layered cost drivers. High-level inputs characterize mission requirements, system performance, and relevant economic factors. Design alternatives are extracted from a standard, product-specific work breakdown structure to pre-load lower-level cost driver inputs and generate the cost-risk analysis. As product design progresses and matures the lower level more detailed cost drivers can be re-accessed and the projected variation of input values narrowed, thereby generating a progressively more accurate estimate of cost-risk. Incorporated into the process-based cost model are techniques for decision analysis, specifically, the analytic hierarchy process (AHP) and functional utility analysis. Design alternatives may then be evaluated not just on cost-risk, but also user defined performance and schedule criteria. This implementation of full-trade study support contributes significantly to the realization of the integrated development environment. The process-based cost estimation model generates development and manufacturing cost estimates. The development team plans to expand the manufacturing process base from approximately 80 manufacturing processes to over 250 processes. Operation and support cost modeling is also envisioned. Process-based estimation considers the materials, resources, and processes in establishing cost-risk and rather depending on weight as an input, actually estimates weight along with cost and schedule.

  9. Hyper-X Flight Engine Ground Testing for X-43 Flight Risk Reduction

    NASA Technical Reports Server (NTRS)

    Huebner, Lawrence D.; Rock, Kenneth E.; Ruf, Edward G.; Witte, David W.; Andrews, Earl H., Jr.

    2001-01-01

    Airframe-integrated scramjet engine testing has been completed at Mach 7 flight conditions in the NASA Langley 8-Foot High Temperature Tunnel as part of the NASA Hyper-X program. This test provided engine performance and operability data, as well as design and database verification, for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe-integrated scramjet data in flight. The Hyper-X Flight Engine, a duplicate Mach 7 X-43 scramjet engine, was mounted on an airframe structure that duplicated the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle trailing edge. This model was also tested to verify and validate the complete flight-like engine system. This paper describes the subsystems that were subjected to flight-like conditions and presents supporting data. The results from this test help to reduce risk for the Mach 7 flights of the X-43.

  10. A conceptual design of shock-eliminating clover combustor for large scale scramjet engine

    NASA Astrophysics Data System (ADS)

    Sun, Ming-bo; Zhao, Yu-xin; Zhao, Guo-yan; Liu, Yuan

    2017-01-01

    A new concept of shock-eliminating clover combustor is proposed for large scale scramjet engine to fulfill the requirements of fuel penetration, total pressure recovery and cooling. To generate the circular-to-clover transition shape of the combustor, the streamline tracing technique is used based on an axisymmetric expansion parent flowfield calculated using the method of characteristics. The combustor is examined using inviscid and viscous numerical simulations and a pure circular shape is calculated for comparison. The results showed that the combustor avoids the shock wave generation and produces low total pressure losses in a wide range of flight condition with various Mach number. The flameholding device for this combustor is briefly discussed.

  11. Enhancement of the CAVE computer code. [aerodynamic heating package for nose cones and scramjet engine sidewalls

    NASA Technical Reports Server (NTRS)

    Rathjen, K. A.; Burk, H. O.

    1983-01-01

    The computer code CAVE (Conduction Analysis via Eigenvalues) is a convenient and efficient computer code for predicting two dimensional temperature histories within thermal protection systems for hypersonic vehicles. The capabilities of CAVE were enhanced by incorporation of the following features into the code: real gas effects in the aerodynamic heating predictions, geometry and aerodynamic heating package for analyses of cone shaped bodies, input option to change from laminar to turbulent heating predictions on leading edges, modification to account for reduction in adiabatic wall temperature with increase in leading sweep, geometry package for two dimensional scramjet engine sidewall, with an option for heat transfer to external and internal surfaces, print out modification to provide tables of select temperatures for plotting and storage, and modifications to the radiation calculation procedure to eliminate temperature oscillations induced by high heating rates. These new features are described.

  12. Scramjet Engine Flowpath Development for the Hyper-X Mach 10 Flight Test

    NASA Technical Reports Server (NTRS)

    Rogers, R. C.; Shih, A. T.; Hass, N. E.

    2005-01-01

    The third flight of the Hyper-X Research Vehicle at nearly Mach 10 in November 2004 demonstrated the potential for airbreathing propulsion at hypersonic speeds. The flowpath of the scramjet engine used in the Hyper-X Research Vehicle was developed using computational fluid dynamic based design method in combination with a systematic series of ground tests in the NASA HyPulse Shock Tunnel at conditions duplicating Mach 10 flight. The test model was a true scale height and length of the internal flight engine flowpath that was partial width and truncated fore and aft. Tests were conducted using a mixture of hydrogen augmented with two-percent silane. Silane gas served as an ignition aid during the short duration of the pulse facility tests. The engine inflow values of pressure, temperature, and Mach number were parametrically varied during the ground test entries to broaden the database over the expected uncertainty in the flight conditions. A review of the ground test technique and comparisons of the ground test pressures along with selected data from the third flight are presented.

  13. Bibliography on propulsion airframe integration technologies for high-speed civil transport applications, 1980-1991

    NASA Technical Reports Server (NTRS)

    Anderson, David J.; Mizukami, Masashi

    1993-01-01

    NASA has initiated the High Speed Research (HSR) program with the goal to develop technologies for a new generation, economically viable, environmentally acceptable, supersonic transport (SST) called the High Speed Civil Transport (HSCT). A significant part of this effort is expected to be in multidisciplinary systems integration, such as in propulsion airframe integration (PAI). In order to assimilate the knowledge database on PAI for SST type aircraft, a bibliography on this subject was compiled. The bibliography with over 1200 entries, full abstracts, and indexes. Related topics are also covered, such as the following: engine inlets, engine cycles, nozzles, existing supersonic cruise aircraft, noise issues, computational fluid dynamics, aerodynamics, and external interference. All identified documents from 1980 through early 1991 are included; this covers the latter part of the NASA Supersonic Cruise Research (SCR) program and the beginnings of the HSR program. In addition, some pre-1980 documents of significant merit or reference value are also included. The references were retrieved via a computerized literature search using the NASA RECON database system.

  14. Suppression of Combustor-Inlet Interaction in a Scramjet Engine under Mach 4 Flight Conditions

    NASA Astrophysics Data System (ADS)

    Kobayashi, Kan; Kanda, Takeshi; Tomioka, Sadatake; Tani, Kouichiro; Sakuranaka, Noboru; Mitani, Tohru

    A sidewall-compression-type scramjet engine was tested under Mach 4 flight conditions. The tested engine had an inlet, a constant cross-sectional area isolator, a constant cross-sectional area combustor, a diverging combustor, and an internal nozzle. In a previous study under the same flight conditions, the maximum thrust increment using fuel injection within the constant-area combustor was 1,380 N at an equivalence ratio of 0.31, and further fuel injection resulted in combustor-inlet interaction (designated as CII). To suppress the CII in the present study, we attempted (1) two-stage fuel injection within the constant-area combustor and the diverging combustor and (2) a boundary layer bleed on the top wall. The former was to suppress heat release around the first-stage fuel injectors in the constant-area combustor, and the latter was to decrease interaction length by decreasing the boundary layer thickness on the top wall. In the case of two-stage fuel injection, the maximum thrust increment was 2,230 N at an equivalence ratio of 0.63. In the case of the boundary layer bleed, on the other hand, the maximum thrust increment was 2,300 N at an equivalence ratio of 0.66. Thus, two-stage fuel injection and boundary layer bleed led to 62% and 67% higher maximum thrust increments than that obtained in the previous study, respectively. Finally, both the two-stage fuel injection and boundary layer bleed were applied simultaneously to obtain the best thrust performance, and the maximum thrust increment was 2,560 N at an equivalence ratio of 0.95. As a result, we obtained an 86% higher maximum thrust increment than that in the previous study. The thrust achievement factor, which was defined as the ratio of the maximum thrusts obtained from experiment and theoretical prediction, under this condition was estimated as 70%.

  15. Comparative Flow Path Analysis and Design Assessment of an Axisymmetric Hydrogen Fueled Scramjet Flight Test Engine at a Mach Number of 6.5

    NASA Technical Reports Server (NTRS)

    McClinton, C.; Rondakov, A.; Semenov, V.; Kopehenov, V.

    1991-01-01

    NASA has contracted with the Central Institute of Aviation Motors CIAM to perform a flight test and ground test and provide a scramjet engine for ground test in the United States. The objective of this contract is to obtain ground to flight correlation for a supersonic combustion ramjet (scramjet) engine operating point at a Mach number of 6.5. This paper presents results from a flow path performance and thermal evaluation performed on the design proposed by the CIAM. This study shows that the engine will perform in the scramjet mode for stoichiometric operation at a flight Mach number of 6.5. Thermal assessment of the structure indicates that the combustor cooling liner will provide adequate cooling for a Mach number of 6.5 test condition and that optional material proposed by CIAM for the cowl leading-edge design are required to allow operation with or without a type IV shock-shock interaction.

  16. OVRhyp, Scramjet Test Aircraft

    NASA Technical Reports Server (NTRS)

    Aslan, J.; Bisard, T.; Dallinga, S.; Draper, K.; Hufford, G.; Peters, W.; Rogers, J.

    1990-01-01

    A preliminary design for an unmanned hypersonic research vehicle to test scramjet engines is presented. The aircraft will be launched from a carrier aircraft at an altitude of 40,000 feet at Mach 0.8. The vehicle will then accelerate to Mach 6 at an altitude of 100,000 feet. At this stage the prototype scramjet will be employed to accelerate the vehicle to Mach 10 and maintain Mach 10 flight for 2 minutes. The aircraft will then decelerate and safely land.

  17. Propellant feed system of a regeneratively cooled scramjet

    SciTech Connect

    Kanda, Takeshi; Masuya, Goro; Wakamatsu, Yoshio )

    1991-04-01

    An expander cycle for an airframe-integrated hydrogen-fueled scramjet is analyzed to study regenerative cooling characteristics and overall specific impulse. Below Mach 10, the specific impulse and thrust coincide with the reference values. At Mach numbers above 10, a reduction of the specific impulse occurs due to the coolant flow rate requirement, which is accompanied by an increase of thrust. It is shown that the thrust may be increased by injecting excess fuel into the combustor to compensate for the decrease of the specific impulse. 9 refs.

  18. A Chief Engineer's View of the NASA X-43A Scramjet Flight Test

    NASA Technical Reports Server (NTRS)

    Marshall, Laurie A.; Corpening, Griffin P.; Sherrill, Robert

    2005-01-01

    This paper presents an overview of the preparation and execution of the first two flights of the NASA X-43A scramjet flight test project. The project consisted of three flights, two planned for Mach 7 and one for Mach 10. The first flight, conducted on June 2, 2001, was unsuccessful and resulted in a nine-month mishap investigation. A two-year return to flight effort ensued and concluded when the second Mach 7 flight was successfully conducted on March 27, 2004. The challenges faced by the project team as they prepared the first ever scramjet-powered airplane for flight are presented. Modifications made to the second flight vehicle as a result of the first flight failure and the return to flight activities are discussed. Flight results and lessons learned are also presented.

  19. CFD Transient Simulation of an Isolator Shock Train in a Scramjet Engine

    DTIC Science & Technology

    2012-09-01

    isolator to contain the PCST is one of two methods to measure isolator performance. The other is to measure the uniformity of the flow exiting the...an important parameter. As a wall property, skin friction can be measured experimentally without complicated, weighty, or bulky flow diagnostics...Krauss. “Daul-Mode Scramjet Operation at a Mach 5 Flight Enthalpy in a Clean Air Test Facility,” Rarefied Gas Dynamics: 24th International

  20. Propulsion airframe integration session overview and review of Lewis PAI efforts

    NASA Technical Reports Server (NTRS)

    Batterton, Peter G.

    1992-01-01

    Propulsion/Airframe Integration (PAI) is a key issue for the High Speed Civil Transport. The aircraft performance, economics, and environmental acceptability can be adversely affected if the integration of the propulsion system and the airframe is not addressed properly or in a timely manner. Some of the goals are listed in this figure. In particular, these goals are highly influenced by how successfully the propulsion system and airframe are integrated. These goals were grouped by the 'Aero' and 'Propulsion' categories to suggest which group of technologists will likely be addressing them. In terms of the NASA High Speed Research Program, the ultimate objective for propulsion/airframe integration is to demonstrate the technologies for achievement of these goals on a 'single' integrated configuration.

  1. A highly efficient engineering tool for three-dimensional scramjet flowfield and heat transfer computations

    NASA Technical Reports Server (NTRS)

    Kamath, Pradeep S.; Hawkins, Richard W.; Baker, Nathaniel R.; Mcclinton, Charles R.

    1991-01-01

    The SIMPLE-based parabolic flow code, SHIP3D, was under development for use as a parametric design and analysis tool for scramjets. Some capabilities and applications of the code are demonstrated, and a report on its current status is given. The focus is on the combustor for which the code was mostly used. Recently, it was also applied to nozzle flows. Code validation results are presented for combustor unit problems involving film cooling, transverse fuel injection, and nozzle test. A parametric study of a film cooled or transpiration cooled Mach 16 combustor is also conducted to illustrate the application of the code to a design problem.

  2. Overview of NATO Background on Scramjet Technology. Chapter 1

    NASA Technical Reports Server (NTRS)

    Drummond, J. Philip; Bouchez, Marc; McClinton, Charles R.

    2006-01-01

    The purpose of the present overview is to summarize the current knowledge of the NATO contributors. All the topics will be addressed in this chapter, with references and some examples. This background enhances the level of knowledge of the NATO scramjet community, which will be used for writing the specific chapters of the Report. Some previous overviews have been published on scramjet technology worldwide. NASA, DOD, the U.S. industry and global community have studied scramjet-powered hypersonic vehicles for over 40 years. Within the U.S. alone, NASA, DOD (DARPA, U.S. Navy and USAF), and industry have participated in hypersonic technology development. Over this time NASA Langley Research Center continuously studied hypersonic system design, aerothermodynamics, scramjet propulsion, propulsion-airframe integration, high temperature materials and structural architectures, and associated facilities, instrumentation and test methods. These modestly funded programs were substantially augmented during the National Aero-Space Plane (X-30) Program, which spent more than $3B between 1984 and 1995, and brought the DOD and other NASA Centers, universities and industry back into hypersonics. In addition, significant progress was achieved in all technologies required for hypersonic flight, and much of that technology was transferred into other programs, such as X-33, DC-X, X-37, X-43, etc. In addition, technology transfer impacted numerous other industries, including automotive, medical, sports and aerospace.

  3. Pioneering scramjet developments by Antonio Ferri

    NASA Technical Reports Server (NTRS)

    Erdos, John I.; Nucci, Louis M.

    1992-01-01

    The concept is summarized of a diffusive burning supersonic combustion ramjet engine (scramjet) envisioned by Antonio Ferri and some of the salient technologies are highlighted as developed by General Applied Science Labs, PIBAL, and NYU, under his direction.

  4. Propulsion Airframe Integration Test Techniques for Hypersonic Airbreathing Configurations at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Witte, David W.; Huebner, Lawrence D.; Trexler, Carl A.; Cabell, Karen F.; Andrews, Earl H., Jr.

    2003-01-01

    The scope and significance of propulsion airframe integration (PAI) for hypersonic airbreathing vehicles is presented through a discussion of the PAI test techniques utilized at NASA Langley Research Center. Four primary types of PAI model tests utilized at NASA Langley for hypersonic airbreathing vehicles are discussed. The four types of PAI test models examined are the forebody/inlet test model, the partial-width/truncated propulsion flowpath test model, the powered exhaust simulation test model, and the full-length/width propulsion flowpath test model. The test technique for each of these four types of PAI test models is described, and the relevant PAI issues addressed by each test technique are illustrated through the presentation of recent PAI test data.

  5. Computational Fluid Dynamics (CFD) Image of Hyper-X Research Vehicle at Mach 7 with Engine Operating

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This computational fluid dynamics (CFD) image shows the Hyper-X vehicle at a Mach 7 test condition with the engine operating. The solution includes both internal (scramjet engine) and external flow fields, including the interaction between the engine exhaust and vehicle aerodynamics. The image illustrates surface heat transfer on the vehicle surface (red is highest heating) and flowfield contours at local Mach number. The last contour illustrates the engine exhaust plume shape. This solution approach is one method of predicting the vehicle performance, and the best method for determination of vehicle structural, pressure and thermal design loads. The Hyper-X program is an ambitious series of experimental flights to expand the boundaries of high-speed aeronautics and develop new technologies for space access. When the first of three aircraft flies, it will be the first time a non-rocket engine has powered a vehicle in flight at hypersonic speeds--speeds above Mach 5, equivalent to about one mile per second or approximately 3,600 miles per hour at sea level. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly

  6. Application of CART3D to Complex Propulsion-Airframe Integration with Vehicle Sketch Pad

    NASA Technical Reports Server (NTRS)

    Hahn, Andrew S.

    2012-01-01

    Vehicle Sketch Pad (VSP) is an easy-to-use modeler used to generate aircraft geometries for use in conceptual design and analysis. It has been used in the past to generate metageometries for aerodynamic analyses ranging from handbook methods to Navier-Stokes computational fluid dynamics (CFD). As desirable as it is to bring high order analyses, such as CFD, into the conceptual design process, this has been difficult and time consuming in practice due to the manual nature of both surface and volume grid generation. Over the last couple of years, VSP has had a major upgrade of its surface triangulation and export capability. This has enhanced its ability to work with Cart3D, an inviscid, three dimensional fluid flow toolset. The combination of VSP and Cart3D allows performing inviscid CFD on complex geometries with relatively high productivity. This paper will illustrate the use of VSP with Cart3D through an example case of a complex propulsion-airframe integration (PAI) of an over-wing nacelle (OWN) airliner configuration.

  7. Turbine Powered Simulator Calibration and Testing for Hybrid Wing Body Powered Airframe Integration

    NASA Technical Reports Server (NTRS)

    Shea, Patrick R.; Flamm, Jeffrey D.; Long, Kurtis R.; James, Kevin D.; Tompkins, Daniel M.; Beyar, Michael D.

    2016-01-01

    Propulsion airframe integration testing on a 5.75% scale hybrid wing body model us- ing turbine powered simulators was completed at the National Full-Scale Aerodynamics Complex 40- by 80-foot test section. Four rear control surface con gurations including a no control surface de ection con guration were tested with the turbine powered simulator units to investigate how the jet exhaust in uenced the control surface performance as re- lated to the resultant forces and moments on the model. Compared to ow-through nacelle testing on the same hybrid wing body model, the control surface e ectiveness was found to increase with the turbine powered simulator units operating. This was true for pitching moment, lift, and drag although pitching moment was the parameter of greatest interest for this project. With the turbine powered simulator units operating, the model pitching moment was seen to increase when compared to the ow-through nacelle con guration indicating that the center elevon and vertical tail control authority increased with the jet exhaust from the turbine powered simulator units.

  8. Subsonic longitudinal aerodynamic characteristics and engine pressure distributions for an aircraft with an integrated scramjet designed for Mach 6 cruise. [conducted in Langley 7 by 10 foot high speed tunnel

    NASA Technical Reports Server (NTRS)

    Huffman, J. K.; Fox, C. H., Jr.; Johnston, P. J.

    1977-01-01

    A 1/10-scale model of a proposed hypersonic aircraft with an integrated scramjet was tested. The investigation took place over a Mach number range from 0.2 to 0.7 and an angle of attack range from 2 deg to approximately 17 deg at a sideslip angle of 0 deg. The primary configuration variables studied were engine location, internal engine geometry, and external engine geometry. The results are presented without analysis.

  9. Technology Roadmap for Dual-Mode Scramjet Propulsion to Support Space-Access Vision Vehicle Development

    NASA Technical Reports Server (NTRS)

    Cockrell, Charles E., Jr.; Auslender, Aaron H.; Guy, R. Wayne; McClinton, Charles R.; Welch, Sharon S.

    2002-01-01

    Third-generation reusable launch vehicle (RLV) systems are envisioned that utilize airbreathing and combined-cycle propulsion to take advantage of potential performance benefits over conventional rocket propulsion and address goals of reducing the cost and enhancing the safety of systems to reach earth orbit. The dual-mode scramjet (DMSJ) forms the core of combined-cycle or combination-cycle propulsion systems for single-stage-to-orbit (SSTO) vehicles and provides most of the orbital ascent energy. These concepts are also relevant to two-stage-to-orbit (TSTO) systems with an airbreathing first or second stage. Foundation technology investments in scramjet propulsion are driven by the goal to develop efficient Mach 3-15 concepts with sufficient performance and operability to meet operational system goals. A brief historical review of NASA scramjet development is presented along with a summary of current technology efforts and a proposed roadmap. The technology addresses hydrogen-fueled combustor development, hypervelocity scramjets, multi-speed flowpath performance and operability, propulsion-airframe integration, and analysis and diagnostic tools.

  10. Shock Tunnel Studies of Scramjet Phenomena 1994

    NASA Technical Reports Server (NTRS)

    Morgan, R. G.; Paull, A.; Stalker, R. J.

    1997-01-01

    Reports by the research staff and graduate students of the Mechanical Engineering Department at the University of Queensland are collected and presented. These reports cover various studies related to the advancement of scramjet technology and the operation of advanced hypervelocity shock-expansion tubes. The report topics include the experimental studies of mixing and combustion in a scramjet flow path, the measurement of integrated thrust and skin friction, and the development of a free-piston-driven expansion tunnel capable of delivering a test gas at superorbital velocities.

  11. Scramjet Combustion Processes

    DTIC Science & Technology

    2010-09-01

    plan for these flights is as follows: Scramjet Combustion Processes RTO-EN-AVT-185 11 - 21 HyShot 5 – A Free-Flying Hypersonic Glider HyShot...5 will be a hypersonic glider designed to fly at Mach 8. It will separate from its rocket booster in space and perform controlled manoeuvres as it...RTO-EN-AVT-185 11 - 1 Scramjet Combustion Processes Michael Smart and Ray Stalker Centre for Hypersonics The University of Queensland

  12. Scramjet Isolator Modeling and Control

    DTIC Science & Technology

    2011-12-01

    Engineering and Management AFIT/DS/ENY/11-19 Abstract The scramjet isolator is a duct in which pressure increases from the inlet to the combustor via a shock... combustor , and the exhaust nozzle which can be seen in Figure 1.1. In the inlet, the flow is initially slowed across a series of shocks prior to...entering the isolator. The isolator allows the flow pressure to increase to that of the combustor . This pressure increase is accomplished across a series of

  13. An improved numerical procedure for the parametric optimization of three dimensional scramjet nozzles. [supersonic combustion ramjet engines - computer programs

    NASA Technical Reports Server (NTRS)

    Dash, S.; Delguidice, P. D.

    1975-01-01

    A parametric numerical procedure permitting the rapid determination of the performance of a class of scramjet nozzle configurations is presented. The geometric complexity of these configurations ruled out attempts to employ conventional nozzle design procedures. The numerical program developed permitted the parametric variation of cowl length, turning angles on the cowl and vehicle undersurface and lateral expansion, and was subject to fixed constraints such as the vehicle length and nozzle exit height. The program required uniform initial conditions at the burner exit station and yielded the location of all predominant wave zones, accounting for lateral expansion effects. In addition, the program yielded the detailed pressure distribution on the cowl, vehicle undersurface and fences, if any, and calculated the nozzle thrust, lift and pitching moments.

  14. Hyper-X Engine Design and Ground Test Program

    NASA Technical Reports Server (NTRS)

    Voland, R. T.; Rock, K. E.; Huebner, L. D.; Witte, D. W.; Fischer, K. E.; McClinton, C. R.

    1998-01-01

    The Hyper-X Program, NASA's focused hypersonic technology program jointly run by NASA Langley and Dryden, is designed to move hypersonic, air-breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated supersonic combustion ramjet propulsion system in flight, providing the first flight validation of wind tunnel, numerical and analytical methods used for design of these vehicles. A substantial portion of the integrated vehicle/engine flowpath development, engine systems verification and validation and flight test risk reduction efforts are experimentally based, including vehicle aeropropulsive force and moment database generation for flight control law development, and integrated vehicle/engine performance validation. The Mach 7 engine flowpath development tests have been completed, and effort is now shifting to engine controls, systems and performance verification and validation tests, as well as, additional flight test risk reduction tests. The engine wind tunnel tests required for these efforts range from tests of partial width engines in both small and large scramjet test facilities, to tests of the full flight engine on a vehicle simulator and tests of a complete flight vehicle in the Langley 8-Ft. High Temperature Tunnel. These tests will begin in the summer of 1998 and continue through 1999. The first flight test is planned for early 2000.

  15. X-51A Scramjet Demonstrator Program: Waverider Ground and Flight Test

    DTIC Science & Technology

    2013-11-01

    USAF) WaveRider program. The overall test objective of the X-51A program was to demonstrate a scramjet engine using endothermic hydrocarbon fuel...Hypersonic Technology (HyTech) scramjet engine, integrated into the vehicle, used endothermic hydrocarbon fuel (JP-7). The X-51A was designed to be

  16. Progress in Applying Tunable Diode Laser Absorption Spectroscopy to Scramjet Isolators and Combustors

    DTIC Science & Technology

    2010-05-01

    AFRL-RZ-WP-TP-2010-2146 PROGRESS IN APPLYING TUNABLE DIODE LASER ABSORPTION SPECTROSCOPY TO SCRAMJET ISOLATORS AND COMBUSTORS Michael S... COMBUSTORS 5a. CONTRACT NUMBER IN HOUSE 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 61102F 6. AUTHOR(S) Michael S. Brown, Skip Williams...paths -- scramjet engines in particular. In this report we summarize our progress in applying TDLAS to scramjet isolators and combustors . 15

  17. Validation of scramjet exhaust simulation technique at Mach 6

    NASA Technical Reports Server (NTRS)

    Hopkins, H. B.; Konopka, W.; Leng, J.

    1979-01-01

    Current design philosophy for hydrogen-fueled, scramjet-powered hypersonic aircraft results in configurations with strong couplings between the engine plume and vehicle aerodynamics. The experimental verification of the scramjet exhaust simulation is described. The scramjet exhaust was reproduced for the Mach 6 flight condition by the detonation tube simulator. The exhaust flow pressure profiles, and to a large extent the heat transfer rate profiles, were then duplicated by cool gas mixtures of Argon and Freon 13B1 or Freon 12. The results of these experiments indicate that a cool gas simulation of the hot scramjet exhaust is a viable simulation technique except for phenomena which are dependent on the wall temperature relative to flow temperature.

  18. Numerical Investigation of the Mixing Process in Inlet-fuelled Scramjets

    NASA Astrophysics Data System (ADS)

    Gehre, R. M.; Peterson, D.; Wheatley, V.; Boyce, R. R.

    For decades, researchers have pursued the development of supersonic combustion ramjet (scramjet) engines. The scramjet is an airbreathing hypersonic propulsion system that utilizes atmospheric oxygen for combustion. Hence, the oxidizer does not have to be carried on-board, which makes scramjets more efficient than rockets and thus very attractive for hypersonic transportation or access-to-space systems. However, hypersonic speeds cause, amongst other challenges, very short residence times within the scramjet combustor. Therefore, achieving a high combustion efficiency poses a major challenge. The combustion efficiency is closely coupled with the mixing efficiency, since most scramjets are mixing limited. Thus, extensive studies focusing on mixing enhancement have been conducted in the past. Different injector shapes and angles [1, 2], flame holding devices [3] and mixing enhancing hypermixers [4] have been investigated in scramjets.

  19. X-43A: The First Flight of a Scramjet Powered Airplane

    NASA Technical Reports Server (NTRS)

    Corpening, Griff

    2004-01-01

    A viewgraph presentation describing the X-43A Scramjet engine is shown. The topics include: 1) Scramjets; 2) Overview of X-43A; 3) What Happened the 1st Time; 4) Return to Flight; and 5) What Happened the 2nd Time.

  20. Hydrogen scramjet with side wall injection

    NASA Technical Reports Server (NTRS)

    Morgan, R. G.; Paull, A.; Morris, N.; Stalker, R. J.

    1986-01-01

    Preliminary results are presented from an experimental evaluation of a hydrogen-burning wall-injection scramjet engine, performed in the free-piston shock tunnel at the Australian National University. The advantages of scramjet propulsion for high-Mach-number high-altitude flight are reviewed, and the need for an alternative to injection struts is indicated. Pressure profiles and heat-transfer measurements for constant-area and diverging ducts are presented graphically and characterized in detail. The ability of the injected flow to shield the chamber wall from the heat of the freestream flow is demonstrated. The extrapolation of the shock-tunnel results to flight conditions, however, is made difficult by the fact that a significant amount of fuel passes through a quenched zone without burning, seriously degrading overall performance.

  1. Shock Driven Multiphase Instabilities in Scramjet Applications

    NASA Astrophysics Data System (ADS)

    McFarland, Jacob

    2016-11-01

    Shock driven multiphase instabilities (SDMI) arise in many applications from dust production in supernovae to ejecta distribution in explosions. At the limit of small, fast reacting particles the instability evolves similar to the Richtmyer-Meshkov (RM) instability. However, as additional particle effects such as lag, phase change, and collisions become significant the required parameter space becomes much larger and the instability deviates significantly from the RM instability. In scramjet engines the SDMI arises during a cold start where liquid fuel droplets are injected and processed by shock and expansion waves. In this case the particle evaporation and mixing is important to starting and sustaining combustion, but the particles are large and slow to react, creating significant multiphase effects. This talk will examine multiphase mixing in scramjet relevant conditions in 3D multiphase hydrodynamic simulations using the FLASH code from the University of Chicago FLASH center.

  2. Sensing and Active Flow Control for Advanced BWB Propulsion-Airframe Integration Concepts

    NASA Technical Reports Server (NTRS)

    Fleming, John; Anderson, Jason; Ng, Wing; Harrison, Neal

    2005-01-01

    In order to realize the substantial performance benefits of serpentine boundary layer ingesting diffusers, this study investigated the use of enabling flow control methods to reduce engine-face flow distortion. Computational methods and novel flow control modeling techniques were utilized that allowed for rapid, accurate analysis of flow control geometries. Results were validated experimentally using the Techsburg Ejector-based wind tunnel facility; this facility is capable of simulating the high-altitude, high subsonic Mach number conditions representative of BWB cruise conditions.

  3. Calculated Effects of Nitric Oxide Flow Contamination on Scramjet Performance

    NASA Technical Reports Server (NTRS)

    Fischer, Karen E.; Rock, Kenneth E.

    1995-01-01

    The level of nitric oxide contamination in the test gas of the NASA Langley Research Center Arc-Heated Scramjet Test Facility and the effect of the contamination on scramjet test engine performance were investigated analytically. The study was conducted for standard facility conditions corresponding to Mach 6, 7, and 8 flight simulations. The analytically determined levels of nitric oxide produced in the facility are compared with experimentally measured levels. Results of the analysis indicate that nitric oxide levels range from one to three mole percent, which corroborates the measured levels. A three-stream combustor code with finite rate chemistry was used to investigate how nitric oxide affects scramjet performance in terms of combustor pressure rise, heat release, and thrust performance. Results indicate minimal effects on engine performance for the test conditions of this investigation.

  4. Shock Tunnel Studies of Scramjet Phenomena 1995. Supplement 13

    NASA Technical Reports Server (NTRS)

    Morgan, R. G.; Stalker, R. J.; Paull, A.

    1997-01-01

    Reports by the research staff and graduate students of the Mechanical Engineering Department at the University of Queensland are collected and presented. These reports cover various studies related to the advancement of scramjet technology and the operation of advanced hypervelocity shock-expansion tubes. The report topics include the experimental studies of mixing and combustion in a scramjet flow path, the measurement of integrated thrust and skin friction, and the development of a free-piston-driven expansion tunnel capable delivering a test gas at super orbital velocities.

  5. Ramjets: Airframe Integration

    DTIC Science & Technology

    2010-09-01

    Thermal Management (Propulsion a vitesse elevee : Conception du moteur - integration et gestion thermique) 14. ABSTRACT These notes deal with the...to those of the ambient air (γ=1.4) • No effects of fuel mass addition (mf << ma) • Complete combustion of the fuel yielding CO2 and H2O as reaction...which leads to a nozzle exit plane static pressure pe that is not necessarily equal to the ambient pressure p0. At zero angle-of-attack the inlet

  6. Ignition problems in scramjet testing

    SciTech Connect

    Mitani, Tohru

    1995-05-01

    Ignition of H{sub 2} in heated air containing H{sub 2}O, radicals, and dust was investigated for scramjet testing. Using a reduced kinetic model for H{sub 2}{minus}O{sub 2} systems, the effects of H{sub 2}O and radicals in nozzles are discussed in relation to engine testing with vitiation heaters. Analysis using linearized rate-equations suggested that the addition of O atoms was 1.5 times more effective than the addition of H atoms for ignition. This result can be applied to the problem of premature ignition caused by residual radicals and to plasma-jet igniters. Thermal and chemical effects of dust, inevitable in storage air heaters, were studied next. The effects of heat capacity and size of dust were expressed in terms of an exponential integral function. It was found that the radical termination on the surface of dust produces an effect equivalent to heat loss. The inhibition of ignition by dust may result, if the mass fraction of dust becomes 10{sup {minus}3}.

  7. Nozzle performance calibration and turbomachinery operational analysis of Turbo-Powered Simulators (TPS) for the NASA-Langley EET propulsion airframe integration investigation

    NASA Technical Reports Server (NTRS)

    Rowe, R. K.

    1982-01-01

    The Boeing flight simulator chamber facility energy efficient transport (EET) propulsion/airframe integration investigation was tested. Two turbopowered simulators (TPS) were calibrated with four different nacelle configurations. Two nacelle configurations were calibrated on each TPS, each calibration simulating the full nozzle pressure ratio range encountered in a one atmosphere total pressure wind tunnel operating over a Mach number range from 0.70 to 0.90 Mach. The results of the calibration test are in the form of velocity and discharge coefficient which is used in determining the simulator thrust while the TPS is operating in the wind tunnel.

  8. Airframe-Jet Engine Integration Noise

    NASA Technical Reports Server (NTRS)

    Tam, Christopher; Antcliff, Richard R. (Technical Monitor)

    2003-01-01

    It has been found experimentally that the noise radiated by a jet mounted under the wing of an aircraft exceeds that of the same jet in a stand-alone environment. The increase in noise is referred to as jet engine airframe integration noise. The objectives of the present investigation are, (1) To obtain a better understanding of the physical mechanisms responsible for jet engine airframe integration noise or installation noise. (2) To develop a prediction model for jet engine airframe integration noise. It is known that jet mixing noise consists of two principal components. They are the noise from the large turbulence structures of the jet flow and the noise from the fine scale turbulence. In this investigation, only the effect of jet engine airframe interaction on the fine scale turbulence noise of a jet is studied. The fine scale turbulence noise is the dominant noise component in the sideline direction. Thus we limit out consideration primarily to the sideline.

  9. Design and multifidelity analysis of dual mode scramjet compression system using coupled NPSS and fluent simulation

    NASA Astrophysics Data System (ADS)

    Vijayakumar, Nandakumar

    Hypersonic airbreathing engines mark a potential future development of the aerospace industry and immense efforts have been taken in gaining knowledge in them for the past decades. The physical phenomenon occurring at the hypersonic flow regime makes the design and performance prediction of a scramjet engine hard. Though cutting-edge simulation tools fight their way toward accurate prediction of the environment, the time consumed by the entire process in designing and analyzing a scramjet engine and its component may be exorbitant. A multi-fidelity approach for designing a scramjet with a cruising Mach number of 6 is detailed in this research where high-order simulations are applied according to the physics involved in the component. Two state-of-the-art simulation tools were used to take the aerodynamic and propulsion disciplines into account for realistic prediction of the individual components as well as the entire scramjet. The specific goal of this research is to create a virtual environment to design and analyze a hypersonic, two-dimensional, planar inlet and isolator to check its operability for a dual-mode scramjet engine. The dual mode scramjet engine starts at a Mach number of 3.5 where it operates as a ramjet and accelerates to Mach 6 to be operated as a scramjet engine. The intercomponent interaction between the compression components with the rest of the engine is studied by varying the fidelity of the numerical simulation according to the complexity of the situation. Efforts have been taken to track the transition Mach number as it switches from ramjet to scramjet. A complete scramjet assembly was built using the Numerical Propulsion Simulation System (NPSS) and the performance of the engine was evaluated for various scenarios. Different numerical techniques were opted for varying the fidelity of the analysis with the highest fidelity consisting of 2D RANS CFD simulation. The interaction between the NPSS elements with the CFD solver is governed by the

  10. Simulation of a hydrocarbon fueled scramjet exhaust

    NASA Technical Reports Server (NTRS)

    Leng, J.

    1982-01-01

    Exhaust nozzle flow fields for a fully integrated, hydrocarbon burning scramjet were calculated for flight conditions of M (undisturbed free stream) = 4 at 6.1 km altitude and M (undisturbed free stream) = 6 at 30.5 km altitude. Equilibrium flow, frozen flow, and finite rate chemistry effects are considered. All flow fields were calculated by method of characteristics. Finite rate chemistry results were evaluated by a one dimensional code (Bittker) using streamtube area distributions extracted from the equilibrium flow field, and compared to very slow artificial rate cases for the same streamtube area distribution. Several candidate substitute gas mixtures, designed to simulate the gas dynamics of the real engine exhaust flow, were examined. Two mixtures are found to give excellent simulations of the specified exhaust flow fields when evaluated by the same method of characteristics computer code.

  11. Development of a Premixed Combustion Capability for Scramjet Combustion Experiments

    NASA Technical Reports Server (NTRS)

    Rockwell, Robert D.; Goyne, Christopher P.; Rice, Brian E.; Chelliah, Harsha; McDaniel, James C.; Edwards, Jack R.; Cantu, Luca M. L.; Gallo, Emanuela C. A.; Cutler, Andrew D.; Danehy, Paul M.

    2015-01-01

    Hypersonic air-breathing engines rely on scramjet combustion processes, which involve high speed, compressible, and highly turbulent flows. The combustion environment and the turbulent flames at the heart of these engines are difficult to simulate and study in the laboratory under well controlled conditions. Typically, wind-tunnel testing is performed that more closely approximates engine testing rather than a careful investigation of the underlying physics that drives the combustion process. The experiments described in this paper, along with companion data sets being developed separately, aim to isolate the chemical kinetic effects from the fuel-air mixing process in a dual-mode scramjet combustion environment. A unique fuel injection approach is taken that produces a nearly uniform fuel-air mixture at the entrance to the combustor. This approach relies on the precombustion shock train upstream of the dual-mode scramjet combustor. A stable ethylene flame anchored on a cavity flameholder with a uniformly mixed combustor inflow has been achieved in these experiments allowing numerous companion studies involving coherent anti-Stokes Raman scattering (CARS), particle image velocimetry (PIV), and planar laser induced fluorescence (PLIF) to be performed.

  12. Navier-Stokes calculations of scramjet-nozzle-afterbody flowfields

    NASA Technical Reports Server (NTRS)

    Baysal, Oktay

    1991-01-01

    A comprehensive computational fluid dynamics effort was conducted from 1987 to 1990 to properly design a nozzle and lower aft end of a generic hypersonic vehicle powered by a scramjet engine. The interference of the exhaust on the control surfaces of the vehicle can have adverse effects on its stability. Two-dimensional Navier-Stokes computations were performed, where the exhaust gas was assumed to be air behaving as a perfect gas. Then the exhaust was simulated by a mixture of Freon-12 and argon, which required solving the Navier-Stokes equations for four species: (nitrogen, oxygen, Freon-12, and argon). This allowed gamma to be a field variable during the mixing of the multispecies gases. Two different mixing models were used and comparisons between them as well as the perfect gas air calculations were made to assess their relative merits. Finally, the three dimensional Navier-Stokes computations were made for the full-span scramjet nozzle afterbody module.

  13. Conceptual model of turbulent flameholding for scramjet combustors

    NASA Technical Reports Server (NTRS)

    Huber, P. W.

    1980-01-01

    New concepts and approaches to scramjet combustor design are presented. Blowoff was from failure of the recirculation-zone (RZ) flame to reach the dividing streamline (DS) at the rear stagnation zone. Increased turbulent exchange across the DS helped flameholding due to forward movement of the flame anchor point inside the RZ. Modeling of the blowoff phenomenon was based on a mass conservation concept involving the traverse of a flame element across the RZ and a flow element along the DS. The scale required to achieve flameholding, predicted by the model, showed a strong adverse effect of low pressure and low fuel equivalence ratio, moderate effect of flight Mach number, and little effect of temperature recovery factor. Possible effects of finite rate chemistry on flameholding and flamespreading in scramjets are discussed and recommendations for approaches to engine combustor design as well as for needed research to reduce uncertainties in the concepts are made.

  14. Infrared measurements of a scramjet exhaust. [to determine combustion efficiency

    NASA Technical Reports Server (NTRS)

    Reed, R. A.; Slack, M. W.

    1980-01-01

    Diagnostic 2 - 5 mm infrared spectra of a hydrogen burning scramjet exhaust were measured with an interferometer spectrometer. Exhaust gas temperatures and water vapor partial pressures were determined from the observed intensity and spectral profile of the H2O 2.7 mm infrared emission band. Overall engine combustion efficiencies were derived by combining these measurements with the known engine operating conditions. Efficiencies fall (70 - 50 percent) as fuel equivalence ratios rise (0.4 - 1.0). Data analysis techniques and sensitivity studies are also presented.

  15. Fuel Cell Airframe Integration Study for Short-Range Aircraft. Volume 1; Aircraft Propulsion and Subsystems Integration Evaluation

    NASA Technical Reports Server (NTRS)

    Gummalla, Mallika; Pandy, Arun; Braun, Robert; Carriere, Thierry; Yamanis, Jean; Vanderspurt, Thomas; Hardin, Larry; Welch, Rick

    2006-01-01

    The objective of this study is to define the functionality and evaluate the propulsion and power system benefits derived from a Solid Oxide Fuel Cell (SOFC) based Auxiliary Power Unit (APU) for a future short range commercial aircraft, and to define the technology gaps to enable such a system. United Technologies Corporation (UTC) Integrated Total Aircraft Power System (ITAPS) methodologies were used to evaluate a baseline aircraft and several SOFC architectures. The technology benefits were captured as reductions of the mission fuel burn, life cycle cost, noise and emissions. As a result of the study, it was recognized that system integration is critical to maximize benefits from the SOFC APU for aircraft application. The mission fuel burn savings for the two SOFC architectures ranged from 4.7 percent for a system with high integration to 6.7 percent for a highly integrated system with certain technological risks. The SOFC APU itself produced zero emissions. The reduction in engine fuel burn achieved with the SOFC systems also resulted in reduced emissions from the engines for both ground operations and in flight. The noise level of the baseline APU with a silencer is 78 dBA, while the SOFC APU produced a lower noise level. It is concluded that a high specific power SOFC system is needed to achieve the benefits identified in this study. Additional areas requiring further development are the processing of the fuel to remove sulfur, either on board or on the ground, and extending the heat sink capability of the fuel to allow greater waste heat recovery, resolve the transient electrical system integration issues, and identification of the impact of the location of the SOFC and its size on the aircraft.

  16. Scramjet isolator modeling and control

    NASA Astrophysics Data System (ADS)

    Hutzel, John Robert

    The scramjet isolator is a duct in which pressure increases from the inlet to the combustor via a shock train. The shock train leading edge (LE) location must be controlled in an operational scramjet. A LE location measurement algorithm, dynamic model, and control algorithm were developed and validated with 500 frame per second (FPS) shadowgraph images in this research. The test apparatus consisted of a direct connect cold-flow high-speed wind tunnel with an adjustable ramp mounted in the tunnel floor. Ramp adjustments changed the tunnel cross-sectional area which changed the tunnel back pressure and LE location. Wall-mounted pressure transducers and a high-speed camera were used for data collection. The LE location measurement algorithm is the first with results validated using 500 FPS shadowgraph images to measure the LE location with root mean square (RMS) errors less than 20% of a duct height, D, although the transducers were separated by 50% of D. The developed and validated dynamic model is the first with error RMS values less than 24% of D. Finally, the first control algorithm capable of controlling the LE location within 50% of D was developed and validated.

  17. Assessing the potential for improved scramjet performance through application of electromagnetic flow control

    NASA Astrophysics Data System (ADS)

    Lindsey, Martin Forrester

    Sustained hypersonic flight using scramjet propulsion is the key technology bridging the gap between turbojets and the exoatmospheric environment where a rocket is required. Recent efforts have focused on electromagnetic (EM) flow control to mitigate the problems of high thermomechanical loads and low propulsion efficiencies associated with scramjet propulsion. This research effort is the first flight-scale, three-dimensional computational analysis of a realistic scramjet to determine how EM flow control can improve scramjet performance. Development of a quasi-one dimensional design tool culminated in the first open source geometry of an entire scramjet flowpath. This geometry was then tested extensively with the Air Force Research Laboratory's three-dimensional Navier-Stokes and EM coupled computational code. As part of improving the model fidelity, a loosely coupled algorithm was developed to incorporate thermochemistry. This resulted in the only open-source model of fuel injection, mixing and combustion in a magnetogasdynamic (MGD) flow controlled engine. In addition, a control volume analysis tool with an electron beam ionization model was presented for the first time in the context of the established computational method used. Local EM flow control within the internal inlet greatly impacted drag forces and wall heat transfer but was only marginally successful in raising the average pressure entering the combustor. The use of an MGD accelerator to locally increase flow momentum was an effective approach to improve flow into the scramjet's isolator. Combustor-based MGD generators proved superior to the inlet generator with respect to power density and overall engine efficiency. MGD acceleration was shown to be ineffective in improving overall performance, with all of the bypass engines having approximately 33% more drag than baseline and none of them achieving a self-powered state.

  18. Hypervelocity Capability of the HYPULSE Shock-Expansion Tunnel for Scramjet Testing

    NASA Technical Reports Server (NTRS)

    Foelsche, Robert O.; Rogers, R. Clayton; Tsai, Ching-Yi; Bakos, Robert J.; Shih, Ann T.

    2001-01-01

    New hypervelocity capabilities for scramjet testing have recently been demonstrated in the HYPULSE Shock-Expansion Tunnel (SET). With NASA's continuing interests in scramjet testing at hypervelocity conditions (Mach 12 and above), a SET nozzle was designed and added to the HYPULSE facility. Results of tests conducted to establish SET operational conditions and facility nozzle calibration are presented and discussed for a Mach 15 (M15) flight enthalpy. The measurements and detailed computational fluid dynamics calculations (CFD) show the nozzle delivers a test gas with sufficiently wide core size to be suitable for free-jet testing of scramjet engine models of similar scale as, those tested in conventional low Mach number blow-down test facilities.

  19. A Finite Rate Chemical Analysis of Nitric Oxide Flow Contamination Effects on Scramjet Performance

    NASA Technical Reports Server (NTRS)

    Cabell, Karen F.; Rock, Kenneth E.

    2003-01-01

    The level of nitric oxide contamination in the test gas of the Langley Research Center Arc-Heated Scramjet Test Facility and the effect of the contamination on scramjet test engine performance were investigated analytically. A finite rate chemical analysis was performed to determine the levels of nitric oxide produced in the facility at conditions corresponding to Mach 6 to 8 flight simulations. Results indicate that nitric oxide levels range from one to three mole percent, corroborating previously obtained measurements. A three-stream combustor code with finite rate chemistry was used to investigate the effects of nitric oxide on scramjet performance. Results indicate that nitric oxide in the test gas causes a small increase in heat release and thrust performance for the test conditions investigated. However, a rate constant uncertainty analysis suggests that the effect of nitric oxide ranges from no net effect, to an increase of about 10 percent in thrust performance.

  20. Emission calculations for a scramjet powered hypersonic transport

    NASA Technical Reports Server (NTRS)

    Lezberg, E. A.

    1973-01-01

    Calculations of exhaust emissions from a scramjet powered hypersonic transport burning hydrogen fuel were performed over a range of Mach numbers of 5 to 12 to provide input data for wake mixing calculations and forecasts of future levels of pollutants in the stratosphere. The calculations were performed utilizing a one-dimensional chemical kinetics computer program for the combustor and exhaust nozzle of a fixed geometry dual-mode scramjet engine. Inlet conditions to the combustor and engine size was based on a vehicle of 227,000 kg (500,000 lb) gross take of weight with engines sized for Mach 8 cruise. Nitric oxide emissions were very high for stoichiometric engine operation but for Mach 6 cruise at reduced equivalence ratio are in the range predicted for an advanced supersonic transport. Combustor designs which utilize fuel staging and rapid expansion to minimize residence time at high combustion temperatures were found to be effective in preventing nitric oxide formation from reaching equilibrium concentrations.

  1. Design of three-dimensional scramjet inlets for hypersonic propulsion

    NASA Technical Reports Server (NTRS)

    Simmons, J. M.; Weidner, E. H.

    1986-01-01

    The paper outlines an approach to the design of three-dimensional inlets for scramjet engines. The basis of the techniques used is the method of streamline tracing through an inviscid axisymmetric flow field. A technique is described for making a smooth change of cross-section shape from rectangular to circular. A feature is the considerable use of computer-graphics to provide a 'user-oriented' procedure which can produce promising design configurations for subsequent analysis with CFD codes. An example is given to demonstrate the capabilities of the design techniques.

  2. Evaluation of a storable fluorine based pilot for scramjets

    NASA Technical Reports Server (NTRS)

    Diskin, G. S.; Northam, G. B.

    1986-01-01

    An ignitor/pilot device for scramjet engines has been tested with C2H4, C2H6, and CH4 gaseous hydrocarbon fuels at Mach 2. The pilot was designed to operate with storable nontoxic propellants rather than hazardous pyrophorics; the pilot fluid is a mixture of ethylene, oxygen, and sulfur hexafluoride. The pilot reduced the minimum temperature at which all three hydrocarbon fuels ignited; a comparison was then conducted between combustion efficiency calculations for these test results and those of pilots using nitrogen in place of sulfur hexafluoride, and using silane-hydrogen (yielding equivalent energy). The nitrogen-substituted pilot exhibited the highest efficiency.

  3. Dual-Mode Scramjet Flameholding Operability Measurements

    NASA Technical Reports Server (NTRS)

    Donohue, James M.

    2012-01-01

    Flameholding measurements were made in two different direct connect combustor facilities that were designed to simulate a cavity flameholder in the flowfield of a hydrocarbon fueled dual-mode scramjet combustor. The presence of a shocktrain upstream of the flameholder has a significant impact on the inlet flow to the combustor and on the flameholding limits. A throttle was installed in the downstream end of the test rigs to provide the needed back-pressurization and decouple the operation of the flameholder from the backpressure formed by heat release and thermal choking, as in a flight engine. Measurements were made primarily with ethylene fuel but a limited number of tests were also performed with heated gaseous JP-7 fuel injection. The flameholding limits were measured by ramping inlet air temperature down until blowout was observed. The tests performed in the United Technologies Research Center (UTRC) facility used a hydrogen fueled vitiated air heater, Mach 2.2 and 3.3 inlet nozzles, a scramjet combustor rig with a 1.666 by 6 inch inlet and a 0.65 inch deep cavity. Mean blowout temperature measured at the baseline condition with ethylene fuel, the Mach 2.2 inlet and a cavity pressure of 21 psia was 1502 oR. Flameholding sensitivity to a variety of parameters was assessed. Blowout temperature was found to be most sensitive to fuel injection location and fuel flowrates and surprisingly insensitive to operating pressure (by varying both back-pressurization and inlet flowrate) and inlet Mach number. Video imaging through both the bottom and side wall windows was collected simultaneously and showed that the flame structure was quite unsteady with significant lateral movements as well as movement upstream of the flameholder. Experiments in the University of Virginia (UVa) test facility used a Mach 2 inlet nozzle with a 1 inch by 1.5 inch exit cross section, an aspect ratio of 1.5 versus 3.6 in the UTRC facility. The UVa facility tests were designed to measure the

  4. Investigation of "6X" Scramjet Inlet Configurations

    NASA Technical Reports Server (NTRS)

    Alter, Stephen J.

    2012-01-01

    This work represents an initial attempt to determine what, if any, issues arise from scaling demonstration supersonic combustion scramjets to a flight scale making the engine a viable candidate for both military weapon and civilian access to space applications. The original vehicle sizes tested and flown to date, were designed to prove a concept. With the proven designs, use of the technology for applications as weapon systems or space flight are only possible at six to ten times the original scale. To determine effects of scaling, computations were performed with hypersonic inlets designed to operate a nominal Mach 4 and Mach 5 conditions that are possible within the eight foot high temperature tunnel at NASA Langley Research Center. The total pressure recovery for these inlets is about 70%, while maintaining self start conditions, and providing operable inflow to combustors. Based on this study, the primary scaling effect detected is the strength of a vortex created along the cowl edge causing adverse boundary layer growth in the inlet.

  5. A predication model for combustion modes of the scramjet-powered aerospace vehicle based on the nonlinear features of the isolator flow field

    NASA Astrophysics Data System (ADS)

    Yang, Qingchun; Wang, Hongxin; Chetehouna, Khaled; Gascoin, Nicolas

    2017-01-01

    The supersonic combustion ramjet (scramjet) engine remains the most promising airbreathing engine cycle for hypersonic flight, particularly the high-performance dual-mode scramjet in the range of flight Mach number from 4 to 7, because it can operates under different combustion modes. Isolator is a very key component of the dual-mode scramjet engine. In this paper, nonlinear characteristics of combustion mode transition is theoretically analyzed. The discontinuous sudden changes of static pressure and Mach number are obtained as the mode transition occurs, which emphasizing the importance of predication and control of combustion modes. In this paper, a predication model of different combustion modes is developed based on these these nonlinear features in the isolator flow field. it can provide a valuable reference for control system design of the scramjet-powered aerospace vehicle.

  6. Scramjet Tests in a Shock Tunnel at Flight Mach 7, 10, and 15 Conditions

    NASA Technical Reports Server (NTRS)

    Rogers, R. C.; Shih, A. T.; Tsai, C.-Y.; Foelsche, R. O.

    2001-01-01

    Tests of the Hyper-X scramjet engine flowpath have been conducted in the HYPULSE shock tunnel at conditions duplicating the stagnation enthalpy at flight Mach 7, 10, and 15. For the tests at Mach 7 and 10 HYPULSE was operated as a reflected-shock tunnel; at the Mach 15 condition, HYPULSE was operated as a shock-expansion tunnel. The test conditions matched the stagnation enthalpy of a scramjet engine on an aerospace vehicle accelerating through the atmosphere along a 1000 psf dynamic pressure trajectory. Test parameter variation included fuel equivalence ratios from lean (0.8) to rich (1.5+); fuel composition from pure hydrogen to mixtures of 2% and 5% silane in hydrogen by volume; and inflow pressure and Mach number made by changing the scramjet model mounting angle in the HYPULSE test chamber. Data sources were wall pressures and heat flux distributions and schlieren and fuel plume imaging in the combustor/nozzle sections. Data are presented for calibration of the facility nozzles and the scramjet engine model. Comparisons of pressure distributions and flowpath streamtube performance estimates are made for the three Mach numbers tested.

  7. HSCT Propulsion Airframe Integration Studies

    NASA Technical Reports Server (NTRS)

    Chaney, Steve

    1999-01-01

    The Lockheed Martin spillage study was a substantial effort and is worthy of a separate paper. However, since a paper was not submitted a few of the most pertinent results have been pulled out and included in this paper. The reader is urged to obtain a copy of the complete Boeing Configuration Aerodynamics final 1995 contract report for the complete Lockheed documentation of the spillage work. The supersonic cruise studies presented here focus on the bifurcated - axisymmetric inlet drag delta. In the process of analyzing this delta several test/CFD data correlation problems arose that lead to a correction of the measured drag delta from 4.6 counts to 3.1 counts. This study also lead to much better understanding of the OVERFLOW gridding and solution process, and to increased accuracy of the force and moment data. Detailed observations of the CFD results lead to the conclusion that the 3.1 count difference between the two inlet types could be reduced to approximately 2 counts, with an absolute lower bound of 1.2 counts due to friction drag and the bifurcated lip bevel.

  8. Shock Tunnel Studies of Scramjet Phenomena 1993

    NASA Technical Reports Server (NTRS)

    Stalker, R. J.; Bakos, R. J.; Morgan, R. G.; Porter, L.; Mee, D.; Paull, A.; Tuttle, S.; Simmons, J. M.; Wendt, M.; Skinner, K.

    1995-01-01

    Reports by the staff of the University of Queensland on various research studies related to the advancement of scramjet technology and hypervelocity pulse test facilities are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation. This research activity is Supplement 10 under NASA Grant NAGw-674.

  9. Numerical Studies on the Performance of Scramjet Combustor with Alternating Wedge-Shaped Strut Injector

    NASA Astrophysics Data System (ADS)

    Choubey, Gautam; Pandey, K. M.

    2017-04-01

    Numerical analysis of the supersonic combustion and flow structure through a scramjet engine at Mach 7 with alternating wedge fuel injection and with three angle of attack (α=-3°, α=0°, α=3°) have been studied in the present research article. The configuration used here is slight modification of the Rabadan et al. scramjet model. Steady two dimensional (2D) Reynolds-averaged Navier-Stokes (RANS) simulation and Shear stress transport (SST) based on k-ω turbulent model is used to predict the shock structure and combustion phenomenon inside the scramjet combustor. All the simulations are done by using Ansys 14-Fluent code. The combustion model used here is the combination of eddy dissipation and finite rate chemistry models since this model avoids Arrhenius calculations in which reaction rates are controlled by turbulence. Present results show that the geometry with negative angle of attack (α=-3°) have lowest ignition delay and it improves the performance of scramjet combustor as compared to geometry with α=0°, α=3°. The combustion phenomena and efficiency is also found to be stronger and highest in case of α=-3°.

  10. Scramjet Fuel Injection Array Optimization Utilizing Mixed Variable Pattern Search With Kriging Surrogates

    DTIC Science & Technology

    2008-03-01

    layer to create effects non -existent at lower speeds [37]. A schematic of a theorized scramjet engine is shown in Figure 1.3 [19]. In this graphic, the...the pressure on inclined bodies based on Newtonian impact theory [17]. The Mach disk center location, empirically calculated in [16], is given by y1 d...to optimize several aerodynamic shapes in hypersonic flow based on Newtonian flow theory. They reported improved convergence over strictly

  11. Hypersonic ramjet experiment project. Phase 1: Computer program description, ramjet and scramjet cycle performance

    NASA Technical Reports Server (NTRS)

    Jackson, R. J.; Wang, T. T.

    1974-01-01

    A computer program was developed to describe the performance of ramjet and scramjet cycles. The program performs one dimensional calculations of the equilibrium, real-gas internal flow properties of the engine. The program can be used for the following: (1) preliminary design calculation and (2) design analysis of internal flow properties corresponding to stipulated flow areas. Only the combustion of hydrogen in air is considered in this case.

  12. Experiments on Cruise Propulsion with a Hydrogen Scramjet

    NASA Technical Reports Server (NTRS)

    Stalker, R. J.; Paull, A.

    1998-01-01

    Measurement of drag have been made in a shock tunnel on a simple integrated vehicle engine combination for hypersonic cruise with hydrogen scramjet propulsion. The test flow Mach number was 6.4, and the velocity was 2.45 kms(exp -1). Zero Drag, which is the necessary condition for cruise, was achieved as the equivalence ratio approached one. It was found that an analysis using established aerodynamic concept was adequate for predicting drag in the case of no combustion. When combustion occurred results of direct connect experiments provided was qualitative guide to the measured levels of drag, and indicated that thrust nozzle combustion was taking place. An heuristic analysis is used to point to the important effect this may have on propulsive lift.

  13. Supersonic mixing and combustion in scramjets

    NASA Technical Reports Server (NTRS)

    Northam, G. B.; Capriotti, D. P.; Byington, C. S.; Greenberg, I.

    1991-01-01

    Experimental and theoretical studies are being conducted to explore techniques to enhance mixing in scramjet combustors using parallel fuel injection from the base of swept and unswept wall-mounted ramps. The experiments reported herein were conducted using Mach 2 and 3 combustor inlet conditions. Supporting computational and cold flow studies indicated that the observed enhanced mixing for the swept ramp configuration is primarily due to the substantially higher degree of vorticity and entrainment generated by the swept trailing edges.

  14. Hypermixer Pylon Fuel Injection for Scramjet Combustors

    DTIC Science & Technology

    2008-09-11

    wall cavities, intrusive geometries like pylons and struts , or a combination thereof. The type of fuel injection strategy largely depends on the...be affixed to combustor walls or placed on intru- sive geometries like struts or pylons. Wall injection comes with an inherent combustor airflow...only one wall . Using in-stream struts and pylons as fueling devices has been a common practice in scramjet design. Much research continues to be

  15. Scramjet Thermal Management (Tenue thermique des superstatoreacteurs)

    DTIC Science & Technology

    2010-09-01

    combustion). Airbreathing launchers could typically use hydrogen-fuelled DMR. Less energetic fuels like hydrocarbons could also be used at a Mach number ...failing to comply with a collection of information if it does not display a currently valid OMB control number . 1. REPORT DATE SEP 2010 2. REPORT TYPE...N/A 3. DATES COVERED - 4. TITLE AND SUBTITLE Scramjet Thermal Management (Tenue thermique des superstatoréacteurs) 5a. CONTRACT NUMBER 5b

  16. Preliminary scramjet design for hypersonic airbreathing missile application

    NASA Technical Reports Server (NTRS)

    Carlson, C. H.

    1983-01-01

    A conceptual design study of a scramjet engine was conducted for a hypersonic surface to air missile (HYSAM). The definition of the engine was based upon the requirements of accelerating the HYSAM from Mach 4 at 20,000 feet to Mach 6 at 100,000 feet and the cruise conditions at Mach 6. The resulting external and internal environmental conditions were used by various engineering disciplines performing design, stress and heat transfer analysis. A detailed structural analysis was conducted along with an indepth thermal analysis. Structurally all the components within the system exhibit positive margins of safety. A feasible concept was defined which uses state-of-the-art materials and existing TMC technology. The engine basically consists of a three dimensional carbon/carbon combustor/nozzle secured to an FS-85 columbium inlet. The carbon/carbon liner is sheathed with carbon felt insulation to thermally protect the FS-85 structure and skin. The thermal analysis of the engine indicates that a thermally viable configuration exists.

  17. Hypersonic Airbreathing Missile

    NASA Technical Reports Server (NTRS)

    Hunt, J. L.; Lawing, P. L.; Marcum, D. C., Jr. (Inventor)

    1978-01-01

    A hypersonic airbreathing missile using dual mode scramjet engines for propulsion is described. The fuselage is constructed of a material with a high heat sink capacity and is covered with a thermal protective shield and lined with an internal insulating blanket. The engine airframe integration uses the flat lower portion of the lower fuselage to precompress the air entering the scramjet engines. The precompression of air entering the scramjet inlets increases as the angles of attack. This feature results in a highly maneuverable missile which can accelerate as it banks into a turn.

  18. Luminescent Measurement Systems for the Investigation of a Scramjet Inlet-Isolator

    PubMed Central

    Idris, Azam Che; Saad, Mohd Rashdan; Zare-Behtash, Hossein; Kontis, Konstantinos

    2014-01-01

    Scramjets have become a main focus of study for many researchers, due to their application as propulsive devices in hypersonic flight. This entails a detailed understanding of the fluid mechanics involved to be able to design and operate these engines with maximum efficiency even at their off-design conditions. It is the objective of the present cold-flow investigation to study and analyse experimentally the mechanics of the fluid structures encountered within a generic scramjet inlet at M = 5. Traditionally, researchers have to rely on stream-thrust analysis, which requires the complex setup of a mass flow meter, a force balance and a heat transducer in order to measure inlet-isolator performance. Alternatively, the pitot rake could be positioned at inlet-isolator exit plane, but this method is intrusive to the flow, and the number of pitot tubes is limited by the model size constraint. Thus, this urgent need for a better flow diagnostics method is addressed in this paper. Pressure-sensitive paint (PSP) has been applied to investigate the flow characteristics on the compression ramp, isolator surface and isolator sidewall. Numerous shock-shock interactions, corner and shoulder separation regions, as well as shock trains were captured by the luminescent system. The performance of the scramjet inlet-isolator has been shown to improve when operated in a modest angle of attack. PMID:24721773

  19. System design and analysis of hydrocarbon scramjet with regeneration cooling and expansion cycle

    NASA Astrophysics Data System (ADS)

    Wu, Xianyu; Yang, Jun; Zhang, Hua; Shen, Chibing

    2015-06-01

    A new expansion cycle scheme of the scramjet engine system including a hydrocarbon-fuel-based (kerosene) regenerativecooling system and supercritical/cracking kerosene-based turbo-pump was proposed in this paper. In this cycle scheme, the supercritical/cracking kerosene with high pressure and high temperature is formed through the cooling channel. And then, in order to make better use of the high energy of the supercritical/cracking fuel, the supercritical/cracking kerosene fuel was used to drive the turbo-pump to obtain a high pressure of the cold kerosenefuel at the entrance of the cooling channel. In the end, the supercritical/cracking kerosene from the turbine exit is injected into the scramjet combustor. Such supercritical/cracking kerosene fuel can decrease the fuel-air mixing length and increase the combustion efficiency, due to the gas state and low molecular weight of the cracking fuel. In order to ignite the cold kerosene in the start-up stage, the ethylene-assisted ignition subsystem was applied. In the present paper, operating modes and characteristics of the expansion cycle system are first described.And then, the overall design of the system and the characterisitics of the start-up process are analyzed numerically to investigate effects of the system parameters on the scramjet start-up performance. The results show that the expansion cycle system proposed in this paper can work well under typical conditions. The research work in this paper can help to make a solid foundation for the research on the coupling characteristics between the dynamicsand thermodynamics of the scramjet expansion cycle system

  20. Shock tunnel studies of scramjet phenomena, supplement 6

    NASA Technical Reports Server (NTRS)

    Wendt, M.; Nettleton, M.; Morgan, R. G.; Skinner, K.; Casey, R.; Stalker, R.; Brescianini, C.; Paull, A.; Allen, G.; Smart, M.

    1993-01-01

    Reports by the staff of the University of Queensland on various research studies related to the advancement of scramjet technology are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation.

  1. Shock tunnel studies of scramjet phenomena, supplement 7

    NASA Technical Reports Server (NTRS)

    Bakos, R. J.; Morgan, R. G.; Tuttle, S. L.; Kelly, G. M.; Paull, A.; Simmons, J. M.; Stalker, R. J.; Pulsonetti, M. V.; Buttsworth, D.; Allen, G. A., Jr.

    1993-01-01

    Reports by the staff of the University of Queensland on various research studies related to the advancement of scramjet technology are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation. This research activity is Supplement 7 under NASA Grant NAGW-674.

  2. Shock tunnel studies of scramjet phenomena, supplement 8

    NASA Technical Reports Server (NTRS)

    Stalker, R. J.; Hollis, P.; Allen, G. A.; Roberts, G. T.; Tuttle, S.; Bakos, R. J.; Morgan, R. G.; Pulsonetti, M. V.; Brescianini, C.; Buttsworth, D. R.

    1993-01-01

    Reports by the staff of the University of Oueensland on various research studies related to the advancement of scramjet technology are presented. These reports document the tests conducted in the reflected shock tunnel T4 and supporting research facilities that have been used to study the injection, mixing, and combustion of hydrogen fuel in generic scramjets at flow conditions typical of hypersonic flight. In addition, topics include the development of instrumentation and measurement technology, such as combustor wall shear and stream composition in pulse facilities, and numerical studies and analyses of the scramjet combustor process and the test facility operation. This research activity is Supplement 8 under NASA Grant NAGW-674.

  3. NASCRIN - NUMERICAL ANALYSIS OF SCRAMJET INLET

    NASA Technical Reports Server (NTRS)

    Kumar, A.

    1994-01-01

    The NASCRIN program was developed for analyzing two-dimensional flow fields in supersonic combustion ramjet (scramjet) inlets. NASCRIN solves the two-dimensional Euler or Navier-Stokes equations in conservative form by an unsplit, explicit, two-step finite-difference method. A more recent explicit-implicit, two-step scheme has also been incorporated in the code for viscous flow analysis. An algebraic, two-layer eddy-viscosity model is used for the turbulent flow calculations. NASCRIN can analyze both inviscid and viscous flows with no struts, one strut, or multiple struts embedded in the flow field. NASCRIN can be used in a quasi-three-dimensional sense for some scramjet inlets under certain simplifying assumptions. Although developed for supersonic internal flow, NASCRIN may be adapted to a variety of other flow problems. In particular, it should be readily adaptable to subsonic inflow with supersonic outflow, supersonic inflow with subsonic outflow, or fully subsonic flow. The NASCRIN program is available for batch execution on the CDC CYBER 203. The vectorized FORTRAN version was developed in 1983. NASCRIN has a central memory requirement of approximately 300K words for a grid size of about 3,000 points.

  4. Shock tunnel studies of scramjet phenomena, supplement 5

    NASA Technical Reports Server (NTRS)

    Casey, R.; Stalker, R. J.; Brescianini, C. P.; Morgan, R. G.; Jacobs, P. A.; Wendt, M.; Ward, N. R.; Akman, N.; Allen, G. A.; Skinner, K.

    1990-01-01

    A series of reports are presented on SCRAMjet studies, shock tunnel studies, and expansion tube studies. The SCRAMjet studies include: (1) Investigation of a Supersonic Combustion Layer; (2) Wall Injected SCRAMjet Experiments; (3) Supersonic Combustion with Transvers, Circular, Wall Jets; (4) Dissociated Test Gas Effects on SCRAMjet Combustors; (5) Use of Silane as a Fuel Additive for Hypersonic Thrust Production, (6) Pressure-length Correlations in Supersonic Combustion; (7) Hot Hydrogen Injection Technique for Shock Tunnels; (8) Heat Release - Wave Interaction Phenomena in Hypersonic Flows; (9) A Study of the Wave Drag in Hypersonic SCRAMjets; (10) Parametric Study of Thrust Production in the Two Dimensional SCRAMjet; (11) The Design of a Mass Spectrometer for use in Hypersonic Impulse Facilities; and (12) Development of a Skin Friction Gauge for use in an Impulse Facility. The shock tunnel studies include: (1) Hypervelocity flow in Axisymmetric Nozzles; (2) Shock Tunnel Development; and (3) Real Gas Efects in Hypervelocity Flows over an Inclined Cone. The expansion tube studies include: (1) Investigation of Flow Characteristics in TQ Expansion Tube; and (2) Disturbances in the Driver Gas of a Shock Tube.

  5. Numerical simulation of scramjet inlet flow fields

    NASA Technical Reports Server (NTRS)

    Kumar, Ajay

    1986-01-01

    A computer program was developed to analyze supersonic combustion ramjet (scramjet) inlet flow fields. The program solves the three-dimensional Euler or Reynolds averaged Navier-Stokes equations in full conservation form by either the fully explicit or explicit-implicit, predictor-corrector method of MacCormack. Turbulence is modeled by an algebraic eddy-viscosity model. The analysis allows inclusion of end effects which can significantly affect the inlet flow field. Detailed laminar and turbulent flow results are presented for a symmetric-wedge corner, and comparisons are made with the available experimental results to allow assessment of the program. Results are then presented for two inlet configurations for which experimental results exist at the NASA Langley Research Center.

  6. Assessment of turbulent models for scramjet flowfields

    NASA Technical Reports Server (NTRS)

    Sindir, M. M.; Harsha, P. T.

    1982-01-01

    The behavior of several turbulence models applied to the prediction of scramjet combustor flows is described. These models include the basic two equation model, the multiple dissipation length scale variant of the two equation model, and the algebraic stress model (ASM). Predictions were made of planar backward facing step flows and axisymmetric sudden expansion flows using each of these approaches. The formulation of each of these models are discussed, and the application of the different approaches to supersonic flows is described. A modified version of the ASM is found to provide the best prediction of the planar backward facing step flow in the region near the recirculation zone, while the basic ASM provides the best results downstream of the recirculation. Aspects of the interaction of numerica modeling and turbulences modeling as they affect the assessment of turbulence models are discussed.

  7. Flow establishment in a generic scramjet combustor

    NASA Astrophysics Data System (ADS)

    Jacobs, P. A.; Rogers, R. C.; Weidner, E. H.; Bittner, R. D.

    1990-10-01

    The establishment of a quasi-steady flow in a generic scramjet combustor was studied for the case of a time varying inflow to the combustor. Such transient flow is characteristic of the reflected shock tunnel and expansion tube test facilities. Several numerical simulations of hypervelocity flow through a straight duct combustor with either a side wall step fuel injector or a centrally located strut injector are presented. Comparisons were made between impulsively started but otherwise constant flow conditions (typical of the expansion tube or tailored operations of the reflected shock tunnel) and the relaxing flow produced by the 'undertailored' operations of the reflected shock tunnel. Generally the inviscid flow features, such as the shock pattern and pressure distribution, were unaffected by the time varying inlet conditions and approached steady state in approx. the times indicated by experimental correlations. However, viscous features, such as heat transfer and skin friction, were altered by the relaxing inlet flow conditions.

  8. Performance Investigation and Characterization of Scramjet and Dual-Mode Scramjet Flow-Fields

    NASA Technical Reports Server (NTRS)

    Riggins, David W.

    2000-01-01

    The following compilation documents significant deliverables under this grant. Note that this summary is extracted from a larger report provided to the Hyper-X office last year at the conclusion of the grant. Current status is documented of the ongoing JANNAF (Joint-Army-Navy-NASA-AirForce) Scramjet Test standards activity from the standpoint of the Analysis SubGroup of which the PI was requested by NASA to be chairman. Also included are some representative contributions to date from the Principle investigator relating to this activity.

  9. Turbulent transport models for scramjet flowfields

    NASA Technical Reports Server (NTRS)

    Sindir, M. M.; Harsha, P. T.

    1984-01-01

    Turbulence modeling approaches were examined from the standpoint of their capability to predict the complex flowfield features observed in scramjet combustions. Thus, for example, the accuracy of each turbulence model, with respect to the prediction of recirculating flows, was examined. It was observed that for large diameter ratio axisymmetric sudden expansion flows, a choice of turbulence model was not critical because of the domination of their flowfields by pressure forces. For low diameter ratio axisymmetric sudden expansions and planar backward-facing steps flows, where turbulent shear stresses are of greater significance, the algebraic Reynolds stress approach, modified to increase its sensitivity to streamline curvature, was found to provide the best results. Results of the study also showed that strongly swirling flows provide a stringent test of turbulence model assumptions. Thus, although flows with very high swirl are not of great practical interest, they are useful for turbulence model development. Finally, it was also noted that numerical flowfields solution techniques have a strong interrelation with turbulence models, particularly with the turbulent transport models which involve source-dominated transport equations.

  10. Further shock tunnel studies of scramjet phenomena

    NASA Technical Reports Server (NTRS)

    Morgan, R. G.; Paull, A.; Morris, N. A.; Stalker, R. J.

    1986-01-01

    Scramjet phenomena were studied using the shock tunnel T3 at the Australian National University. Simple two dimensional models were used with a combination of wall and central injectors. Silane as an additive to hydrogen fuel was studied over a range of temperatures and pressures to evaluate its effect as an ignition aid. The film cooling effect of surface injected hydrogen was measured over a wide range of equivalence. Heat transfer measurements without injection were repeated to confirm previous indications of heating rates lower than simple flat plate predictions for laminar boundary layers in equilibrium flow. The previous results were reproduced and the discrepancies are discussed in terms of the model geometry and departures of the flow from equilibrium. In the thrust producing mode, attempts were made to increase specific impulse with wall injection. Some preliminary tests were also performed on shock induced ignition, to investigate the possibility in flight of injecting fuel upstream of the combustion chamber, where it could mix but not burn.

  11. Fuel-Air Mixing and Combustion in Scramjets

    NASA Technical Reports Server (NTRS)

    Drummond, J. P.; Diskin, Glenn S.; Cutler, A. D.

    2002-01-01

    Activities in the area of scramjet fuel-air mixing and combustion associated with the Research and Technology Organization Working Group on Technologies for Propelled Hypersonic Flight are described. Work discussed in this paper has centered on the design of two basic experiments for studying the mixing and combustion of fuel and air in a scramjet. Simulations were conducted to aid in the design of these experiments. The experimental models were then constructed, and data were collected in the laboratory. Comparison of the data from a coaxial jet mixing experiment and a supersonic combustor experiment with a combustor code were then made and described. This work was conducted by NATO to validate combustion codes currently employed in scramjet design and to aid in the development of improved turbulence and combustion models employed by the codes.

  12. Mach 2 and Mach 3 mixing and combustion in scramjets

    NASA Technical Reports Server (NTRS)

    Northam, G. B.; Capriotti, D. P.; Byington, C. S.; Greenberg, I.

    1991-01-01

    Experimental and theoretical studies are being conducted to explore techniques to enhance mixing in scramjet combustors using parallel fuel injection from the base of swept and unswept wall-mounted ramps. Parallel injection may be useful in high speed scramjets due to the thrust contributed by the momentum of expanding fuel that has been heated in the vehicle cooling cycle. The experiments reported herein were conducted using Mach 2 and 3 combustor inlet conditions. Supporting computational and cold flow studies indicated that the observed enhanced mixing for the swept ramp configuration is primarily due to the substantially higher degree of vorticity and entrainment generated by the swept trailing edges.

  13. Inlet starting predictions for sidewall-compression scramjet inlets

    NASA Technical Reports Server (NTRS)

    Trexler, Carl A.

    1988-01-01

    While inlet capture, total pressure recovery, and compression are important performance parameters in the design of scramjet inlets, the inlet must be capable of starting. An analytical, inviscid method has been developed that provides first order starting information for sidewall-compression scramjet inlets and trends have been established for some of the parameters that affect inlet starting. These parameters include sidewall contraction ratio, cowl position, Mach number, sweep angle, fuel injection struts, and sidewall and strut compression angles. Results from the parametric study are compared with data from both no-strut and two-strut inlet models.

  14. Numerical simulation and modeling of combustion in scramjets

    NASA Astrophysics Data System (ADS)

    Clark, Ryan James

    In the last fifteen years the development of a viable scramjet has quickly approached the following long term goals: responsive sub-orbital space access; long-range, prompt global strike; and high-speed transportation. Nonetheless, there are significant challenges that need to be resolved. These challenges include high skin friction drag and high heat transfer rates, inherent to vehicles in sustained, hypersonic flight. Another challenge is sustaining combustion. Numerical simulation and modeling was performed to provide insight into reducing skin friction drag and sustaining combustion. Numerical simulation was used to investigate boundary layer combustion, which has been shown to reduce skin friction drag. The objective of the numerical simulations was to quantify the effect of fuel injection parameters on boundary layer combustion and ultimately on the change in the skin friction coefficient and heat transfer rate. A qualitative analysis of the results suggest that the reduction in the skin friction coefficient depends on multiple parameters and potentially an interaction between parameters. Sustained combustion can be achieved through a stabilized detonation wave. Additionally, stabilizing a detonation wave will yield rapid combustion. This will allow for a shorter and lighter-weight engine system, resulting in less required combustor cooling. A stabilized detonation wave was numerically modeled for various inlet and geometric cases. The effect of fuel concentration, inlet Mach number, and geometric configuration on the stability of a detonation wave was quantified. Correlations were established between fuel concentration, inlet speed, geometric configuration and parameters characterizing the detonation wave. A linear relationship was quantified between the fuel concentration and the parameters characterizing the detonation wave.

  15. Numerical Simulation of Vitiation Effects on a Hydrogen-Fueled Dual-Mode Scramjet

    NASA Technical Reports Server (NTRS)

    Vyas, Manan A.; Engblom, William A.; Georgiadis, Nicholas J.; Trefny, Charles J.; Bhagwandin, Vishal A.

    2010-01-01

    The Wind-US computational fluid dynamics (CFD) flow solver was used to simulate dual-mode direct-connect ramjet/scramjet engine flowpath tests conducted in the University of Virginia (UVa) Supersonic Combustion Facility (SCF). The objective was to develop a computational capability within Wind-US to aid current hypersonic research and provide insight to flow as well as chemistry details that are not resolved by instruments available. Computational results are compared with experimental data to validate the accuracy of the numerical modeling. These results include two fuel-off non-reacting and eight fuel-on reacting cases with different equivalence ratios, split between one set with a clean (non-vitiated) air supply and the other set with a vitiated air supply (12 percent H2O vapor). The Peters and Rogg hydrogen-air chemical kinetics model was selected for the scramjet simulations. A limited sensitivity study was done to investigate the choice of turbulence model and inviscid flux scheme and led to the selection of the k-epsilon model and Harten, Lax and van Leer (for contact waves) (HLLC) scheme for general use. Simulation results show reasonably good agreement with experimental data and the overall vitiation effects were captured.

  16. A three-dimensional upwind PNS code for chemically reacting scramjet flowfields. [Parabolized Navier Stokes

    SciTech Connect

    Wadawadigi, G.; Tannehill, J.C.; Buelow, P.E.; Lawrence, S.L. NASA, Ames Research Center, Moffett Field, CA )

    1992-07-01

    A new upwind, parabolized Navier-Stokes (PNS) code has been developed to compute the three-dimensional (3D) chemically reacting flow in scramjet (supersonic combustion ramjet) engines. The code is a modification of the 3D upwind PNS (UPS) airflow code which has been extended in the present study to permit internal flow calculations with hydrogen-air chemistry. With these additions, the new code has the capability of computing aerodynamic and propulsive flowfields simultaneously. The algorithm solves the PNS equations using a finite-volume, upwind TVD method based on Roe's approximate Riemann solver that has been modified to account for 'real gas' effects. The fluid medium is assumed to be a chemically reacting mixture of thermally perfect (but calorically imperfect) gases in thermal equilibrium. The new code has been applied to two test cases. These include the Burrows-Kurkov supersonic combustion experiment and a generic 3D scramjet flowfield. The computed results compare favorably with the available experimental data. 38 refs.

  17. Shock Tunnel Studies of Scramjet Phenomena

    NASA Technical Reports Server (NTRS)

    Stalker, R. J.

    1996-01-01

    Work focussed on a large number of preliminary studies of supersonic combustion in a simple combustion duct - thrust nozzle combination, investigating effects of Mach number, equivalence ratio, combustor divergence, fuel injecting angle and other parameters with an influence on the combustion process. This phase lasted for some three or four years, during which strongest emphasis was placed on responding to the request for preliminary experimental information on high enthalpy effects, to support the technology maturation activities of the NASP program. As the need for preliminary data became less urgent, it was possible to conduct more systematic studies of high enthalpy combustion phenomena, and to initiate other projects aimed at improving the facilities and instrumentation used for studying scramjet phenomena at high enthalpies. The combustion studies were particularly directed towards hypersonic combustion, and to the effects of injecting fuel along the combustion chamber wall. A substantial effort was directed towards a study of the effect of scale on the supersonic combustion process. The influence of wave phenomena (both compression waves and expansion waves) on the realization of thrust from a supersonic combustion process was also investigated. The effect of chemical kinetics was looked into, particularly as it affected the composition of the test flow provided by a ground facility. The effect of injection of the fuel through wall orifices was compared with injection from a strut spanning the stream, and the effect of heating the fuel prior to injection was investigated. Studies of fuel-air mixing by shock impingement were also done, as well as mass spectrometer surveys of a combustion wake. The use of hypersonic nozzles with an expansion tube was investigated. A new method was developed for measuring the forces acting of a model in less than one millisecond. Also included in this report are listings of published journal papers and conference presentations.

  18. Free-radicals aided combustion with scramjet applications

    NASA Technical Reports Server (NTRS)

    Yang, Yongsheng; Kumar, Ramohalli

    1992-01-01

    Theoretical and experimental investigations aimed at altering 'nature-prescribed' combustion rates in hydrogen/hydrocarbon reactions with (enriched) air are presented. The intent is to anchor flame zones in supersonic streams, and to ensure proper and controllable complete combustion in scramjets. The diagnostics are nonintrusive through IR thermograms and acoustic emissions in the control and free-radicals altered flame zones.

  19. Analysis of beamed-energy ramjet/scramjet performance

    NASA Technical Reports Server (NTRS)

    Myrabo, L. N.; Powers, M. V.; Zaretzky, C. L.

    1986-01-01

    A study has been performed on a laser-heated ramjet/scramjet vehicle concept for propulsion during the air-breathing portion of an orbital launch trajectory. The concept considers axisymmetric, high-thrust vehicles with external inlets and nozzles. Conceptual design and ramjet/scramjet cycle analysis are emphasized, with propulsive energy provided by combustion of on-board fuel. The conventional ramjet/scramjet combustion chamber is replaced by a laser energy absorption chamber. The elimination of on-board propellant can result in very high thrust-to-weight ratios and payload fractions, in a vehicle with a relatively small degree of mechanical complexity. The basic vehicle has a weight of 12,250 lbf, and a diameter of 5 meters, which is close to the size of the Apollo command module. The ramjet calculations are based on a Mach 3 isentropic inlet with a 13.7 degree half-angle conical tip. The scramjet analysis considers conical inlets with 10, 15, and 30 degree half-angles. Flight Mach numbers from 2 to 20 are considered in the calculations.

  20. The large-amplitude combustion oscillation in a single-side expansion scramjet combustor

    NASA Astrophysics Data System (ADS)

    Ouyang, Hao; Liu, Weidong; Sun, Mingbo

    2015-12-01

    The combustion oscillation in scramjet combustor is believed not existing and ignored for a long time. Compared with the flame pulsation, the large-amplitude combustion oscillation in scramjet combustor is indeed unfamiliar and difficult to be observed. In this study, the specifically designed experiments are carried out to investigate this unusual phenomenon in a single-side expansion scramjet combustor. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The obtained results show that the large-amplitude combustion oscillation can exist in scramjet combustor, which is not occasional and can be reproduced. Under the given conditions of this study, moreover, the large-amplitude combustion oscillation is regular and periodic, whose principal frequency is about 126 Hz. The proceeding of the combustion oscillation is accompanied by the transformation of the flame-holding pattern and combustion mode transition between scramjet mode combustion and ramjet mode combustion.

  1. Three-dimensional shock-shock interactions on the scramjet inlet

    NASA Technical Reports Server (NTRS)

    Singh, D. J.; Tiwari, S. N.; Kumar, A.

    1990-01-01

    The effects of shock impingement on the inlet of a scramjet engine are investigated numerically. The impinging shock is caused by the vehicle forebody. The interaction of this forebody shock with the inlet leading edge shock results in a very complex fully three-dimensional flowfield containing local regions of high pressure and intense heating. In the present investigation, this complex flowfield is calculated by solving the thin-layer Navier-Stokes equations using a finite-volume flux splitting technique due to van Leer. For zero or small sweep angles a Type IV interaction occurs while for moderate sweep of about 25 deg, a Type V interaction occurs. Both Type IV and Type V interactions are investigated.

  2. Three-dimensional shock-shock interactions on the scramjet inlet

    SciTech Connect

    Singh, D.J.; Tiwari, S.N.; Kumar, A.

    1990-01-01

    The effects of shock impingement on the inlet of a scramjet engine are investigated numerically. The impinging shock is caused by the vehicle forebody. The interaction of this forebody shock with the inlet leading edge shock results in a very complex fully three-dimensional flowfield containing local regions of high pressure and intense heating. In the present investigation, this complex flowfield is calculated by solving the thin-layer Navier-Stokes equations using a finite-volume flux splitting technique due to van Leer. For zero or small sweep angles a Type IV interaction occurs while for moderate sweep of about 25 deg, a Type V interaction occurs. Both Type IV and Type V interactions are investigated. 25 refs.

  3. Ramjet/scramjet plus rocket propulsion for a heavy-lift Space Shuttle

    NASA Astrophysics Data System (ADS)

    Lantz, Edward

    1993-10-01

    The possibility of using hydrogen-fueled ramjet/scramjet engines for improving the performance and reducing the operating cost of a second-generation Space Shuttle is examined. For a heavy-lift capability, a two-stage system would be necessary. This could consist of a central Trans Atmospheric Vehicle (TAV) with a hypersonic booster attached to each side. A wheeled ground-based launcher could make the takeoff of such a system possible. By using data from the NASP project and the present Space Shuttle, it is shown that a TAV, which is about 20 percent longer than a Boeing 747, could take a payload of about 200,000 pounds to an earth orbit.

  4. The jet engine design that can drastically reduce oxides of nitrogen

    NASA Technical Reports Server (NTRS)

    Ferri, A.; Agnone, A.

    1974-01-01

    The problem is analyzed for the case of hydrogen fuel, taking into account supersonic and hypersonic vehicles using scramjet engines. The combustion in scramjets occurs at very high velocity and in a short time. In scramjet combustor designs, two different criteria can be used to design the engine. The amount of NO formed in the diffusion flame depends substantially on the maximum temperature reached. Effects of changing the mode of combustion from a diffusion flame to a heat conduction flame are considered, giving attention to the amount of NO produced in an engine of a given design.

  5. Mixing enhancement in a scramjet combustor using fuel jet injection swirl

    NASA Astrophysics Data System (ADS)

    Flesberg, Sonja M.

    The scramjet engine has proven to be a viable means of powering a hypersonic vehicle, especially after successful flights of the X-51 WaveRider and various Hy-SHOT test vehicles. The major challenge associated with operating a scramjet engine is the short residence time of the fuel and oxidizer in the combustor. The fuel and oxidizer have only milliseconds to mix, ignite and combust in the combustion chamber. Combustion cannot occur until the fuel and oxidizer are mixed on a molecular level. Therefore the improvement of mixing is of utmost interest since this can increase combustion efficiency. This study investigated mixing enhancement of fuel and oxidizer within the combustion chamber of a scramjet by introducing swirl to the fuel jet. The investigation was accomplished with numerical simulations using STAR-CCM+ computational fluid dynamic software. The geometry of the University of Virginia Supersonic Combustion Facility was used to model the isolator, combustor and nozzle of a scramjet engine for simulation purposes. Experimental data from previous research at the facility was used to verify the simulation model before investigating the effect of fuel jet swirl on mixing. The model used coaxial fuel jet with a swirling annular jet. Single coaxial fuel jet and dual coaxial fuel jet configurations were simulated for the investigation. The coaxial fuel jets were modelled with a swirling annular jet and non-swirling core jet. Numerical analysis showed that fuel jet swirl not only increased mixing and entrainment of the fuel with the oxidizer but the mixing occurred further upstream than without fuel jet swirl. The burning efficiency was calculated for the all the configurations. An increase in burning efficiency indicated an increase in the mixing of H2 with O2. In the case of the single fuel jet models, the maximum burning efficiency increase due to fuel injection jet swirl was 23.3%. The research also investigated the possibility that interaction between two

  6. Fuel Injector Design Optimization for an Annular Scramjet Geometry

    NASA Technical Reports Server (NTRS)

    Steffen, Christopher J., Jr.

    2003-01-01

    A four-parameter, three-level, central composite experiment design has been used to optimize the configuration of an annular scramjet injector geometry using computational fluid dynamics. The computational fluid dynamic solutions played the role of computer experiments, and response surface methodology was used to capture the simulation results for mixing efficiency and total pressure recovery within the scramjet flowpath. An optimization procedure, based upon the response surface results of mixing efficiency, was used to compare the optimal design configuration against the target efficiency value of 92.5%. The results of three different optimization procedures are presented and all point to the need to look outside the current design space for different injector geometries that can meet or exceed the stated mixing efficiency target.

  7. Numerical Investigation of a Model Scramjet Combustor Using DDES

    NASA Astrophysics Data System (ADS)

    Shin, Junsu; Sung, Hong-Gye

    2017-04-01

    Non-reactive flows moving through a model scramjet were investigated using a delayed detached eddy simulation (DDES), which is a hybrid scheme combining Reynolds averaged Navier-Stokes scheme and a large eddy simulation. The three dimensional Navier-Stokes equations were solved numerically on a structural grid using finite volume methods. An in-house was developed. This code used a monotonic upstream-centered scheme for conservation laws (MUSCL) with an advection upstream splitting method by pressure weight function (AUSMPW+) for space. In addition, a 4th order Runge-Kutta scheme was used with preconditioning for time integration. The geometries and boundary conditions of a scramjet combustor operated by DLR, a German aerospace center, were considered. The profiles of the lower wall pressure and axial velocity obtained from a time-averaged solution were compared with experimental results. Also, the mixing efficiency and total pressure recovery factor were provided in order to inspect the performance of the combustor.

  8. Scramjet Performance Assessment Using Water Absorption Diagnostics (U)

    NASA Technical Reports Server (NTRS)

    Cavolowsky, John A.; Loomis, Mark P.; Deiwert, George

    1995-01-01

    Simultaneous multiple path measurements of temperature and H2O concentration will be presented for the AIMHYE test entries in the NASA Ames 16-Inch Shock Tunnel. Monitoring the progress of high temperature chemical reactions that define scramjet combustor efficiencies is a task uniquely suited to nonintrusive optical diagnostics. One application strategy to overcome the many challenges and limitations of nonintrusive measurements is to use laser absorption spectroscopy coupled with optical fibers. Absorption spectroscopic techniques with rapidly tunable lasers are capable of making simultaneous measurements of mole fraction, temperature, pressure, and velocity. The scramjet water absorption diagnostic was used to measure combustor efficiency and was compared to thrust measurements using a nozzle force balance and integrated nozzle pressures to develop a direct technique for evaluating integrated scramjet performance. Tests were initially performed with a diode laser tuning over a water absorption feature at 1391.7 nm. A second diode laser later became available at a wavelength near 1343.3 nm covering an additional water absorption feature and was incorporated in the system for a two-wavelength technique. Both temperature and mole fraction can be inferred from the lineshape analysis using this approach. Additional high temperature spectroscopy research was conducted to reduce uncertainties in the scramjet application. The lasers are optical fiber coupled to ports at the combustor exit and in the nozzle region. The output from the two diode lasers were combined in a single fiber, and the resultant two-wavelength beam was subsequently split into four legs. Each leg was directed through 60 meters of optical fiber to four combustor exit locations for measurement of beam intensity after absorption by the water within the flow. Absorption results will be compared to 1D combustor analysis using RJPA and nozzle CFD computations as well as to data from a nozzle metric

  9. Study of Magnetogasdynamic Flow Acceleration in a Scramjet Nozzle

    DTIC Science & Technology

    2005-06-01

    direction along the electrodes to be parallel with the flow. At the end of the electrodes, the vectors push the flow toward the wall. The x-y midplane...accelerated case features long regions of slower Mach number along the walls. This is probably due to a combination of the Lorentz forces pushing ...Institute of Tecnology , WPAFB, OH, March 2004. 8. Einstein, A. Relativity, The Special and General Theory . Three Rivers Press, 1961. 9. “Ramjet, Scramjet

  10. Testing Transitions Inside Air-Breathing Scramjet Engines

    NASA Video Gallery

    During their summer internships at NASA centers this year, Aeronautics Academy and Aeronautics Scholarship Program interns produced videos about their work for the NASA Aeronautics "Ideas in Flight...

  11. Aerodynamic Engine/Airframe Integration for High Performance Aircraft and Missiles (L’Integration Aerodynamique des Moteurs et des Cellules dans les Avions et les Missiles a Hautes Performances)

    DTIC Science & Technology

    1992-09-01

    nations. I le conimtent i 4 this putilicmtioii has beccii reproduced direct Ifr hmmi material supplied b\\ A.( iAR) iir thle authors Published September...to move in military A/C development, problem Steady and unsteady distortion of total pressures obtained with hot gas (afterburner) simulation. in...to be made with increased regard to fuel HGR Hot Gas Reingestion economy, environmental noise and first cost; the expanded use M1, MCC) M Local

  12. HIFiRE Direct-Connect Rig (HDCR) Phase I Scramjet Test Results from the NASA Langley Arc-Heated Scramjet Test Facility

    NASA Technical Reports Server (NTRS)

    Cabell, Karen; Hass, Neal; Storch, Andrea; Gruber, Mark

    2011-01-01

    A series of hydrocarbon-fueled direct-connect scramjet ground tests has been completed in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) at simulated Mach 8 flight conditions. These experiments were part of an initial test phase to support Flight 2 of the Hypersonic International Flight Research Experimentation (HIFiRE) Program. In this flight experiment, a hydrocarbon-fueled scramjet is intended to demonstrate transition from dual-mode to scramjet-mode operation and verify the scramjet performance prediction and design tools A performance goal is the achievement of a combusted fuel equivalence ratio greater than 0.7 while in scramjet mode. The ground test rig, designated the HIFiRE Direct Connect Rig (HDCR), is a full-scale, heat sink test article that duplicates both the flowpath lines and a majority of the instrumentation layout of the isolator and combustor portion of the flight test hardware. The primary objectives of the HDCR Phase I tests were to verify the operability of the HIFiRE isolator/combustor across the simulated Mach 6-8 flight regime and to establish a fuel distribution schedule to ensure a successful mode transition. Both of these objectives were achieved prior to the HiFIRE Flight 2 payload Critical Design Review. Mach 8 ground test results are presented in this report, including flowpath surface pressure distributions that demonstrate the operation of the flowpath in scramjet-mode over a small range of test conditions around the nominal Mach 8 simulation, as well as over a range of fuel equivalence ratios. Flowpath analysis using ground test data is presented elsewhere; however, limited comparisons with analytical predictions suggest that both scramjet-mode operation and the combustion performance objective are achieved at Mach 8 conditions.

  13. Advanced stratified charge rotary aircraft engine design study

    NASA Technical Reports Server (NTRS)

    Badgley, P.; Berkowitz, M.; Jones, C.; Myers, D.; Norwood, E.; Pratt, W. B.; Ellis, D. R.; Huggins, G.; Mueller, A.; Hembrey, J. H.

    1982-01-01

    A technology base of new developments which offered potential benefits to a general aviation engine was compiled and ranked. Using design approaches selected from the ranked list, conceptual design studies were performed of an advanced and a highly advanced engine sized to provide 186/250 shaft Kw/HP under cruise conditions at 7620/25,000 m/ft altitude. These are turbocharged, direct-injected stratified charge engines intended for commercial introduction in the early 1990's. The engine descriptive data includes tables, curves, and drawings depicting configuration, performance, weights and sizes, heat rejection, ignition and fuel injection system descriptions, maintenance requirements, and scaling data for varying power. An engine-airframe integration study of the resulting engines in advanced airframes was performed on a comparative basis with current production type engines. The results show airplane performance, costs, noise & installation factors. The rotary-engined airplanes display substantial improvements over the baseline, including 30 to 35% lower fuel usage.

  14. A parametric experimental investigation of a scramjet nozzle at Mach 6 with Freon and argon or air used for exhaust simulation

    NASA Technical Reports Server (NTRS)

    Cubbage, James M.; Monta, William J.

    1991-01-01

    A parametric experimental investigation of a scramjet nozzle was conducted with a gas mixture used to simulate the scramjet engine exhaust flow at a free-stream Reynolds number of approximately 6.5 x 10(exp 6) per foot. External nozzle surface angles of 16, 20, and 24 deg were tested with a fixed-length ramp and for cowl internal surface angles of 6 and 12 deg. Pressure data on the external nozzle surface were obtained for mixtures of Freon and argon gases with a ratio of specific heats of about 1.23, which matches that of a scramjet exhaust. Forces and moments were determined by integration of the pressure data. Two nozzle configurations were also tested with air used to simulate the exhaust flow. On the external nozzle surface, lift and thrust forces for air exhaust simulation were approximately half of those for Freon-argon exhaust simulation and the pitching moment was approximately a third. These differences were primarily due to the difference in the ratios of specific heats between the two exhaust simulation gases. A 20 deg external surface angle produced the greatest thrust for a 6 deg cowl internal surface angle. A flow fence significantly increased lift and thrust forces over those for the nozzle without a flow fence.

  15. Control-Relevant Modeling, Analysis, and Design for Scramjet-Powered Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Rodriguez, Armando A.; Dickeson, Jeffrey J.; Sridharan, Srikanth; Benavides, Jose; Soloway, Don; Kelkar, Atul; Vogel, Jerald M.

    2009-01-01

    Within this paper, control-relevant vehicle design concepts are examined using a widely used 3 DOF (plus flexibility) nonlinear model for the longitudinal dynamics of a generic carrot-shaped scramjet powered hypersonic vehicle. Trade studies associated with vehicle/engine parameters are examined. The impact of parameters on control-relevant static properties (e.g. level-flight trimmable region, trim controls, AOA, thrust margin) and dynamic properties (e.g. instability and right half plane zero associated with flight path angle) are examined. Specific parameters considered include: inlet height, diffuser area ratio, lower forebody compression ramp inclination angle, engine location, center of gravity, and mass. Vehicle optimizations is also examined. Both static and dynamic considerations are addressed. The gap-metric optimized vehicle is obtained to illustrate how this control-centric concept can be used to "reduce" scheduling requirements for the final control system. A classic inner-outer loop control architecture and methodology is used to shed light on how specific vehicle/engine design parameter selections impact control system design. In short, the work represents an important first step toward revealing fundamental tradeoffs and systematically treating control-relevant vehicle design.

  16. An Airbreathing Launch Vehicle Design with Turbine-Based Low-Speed Propulsion and Dual Mode Scramjet High-Speed Propulsion

    NASA Technical Reports Server (NTRS)

    Moses, P. L.; Bouchard, K. A.; Vause, R. F.; Pinckney, S. Z.; Ferlemann, S. M.; Leonard, C. P.; Taylor, L. W., III; Robinson, J. S.; Martin, J. G.; Petley, D. H.

    1999-01-01

    Airbreathing launch vehicles continue to be a subject of great interest in the space access community. In particular, horizontal takeoff and horizontal landing vehicles are attractive with their airplane-like benefits and flexibility for future space launch requirements. The most promising of these concepts involve airframe integrated propulsion systems, in which the external undersurface of the vehicle forms part of the propulsion flowpath. Combining of airframe and engine functions in this manner involves all of the design disciplines interacting at once. Design and optimization of these configurations is a most difficult activity, requiring a multi-discipline process to analytically resolve the numerous interactions among the design variables. This paper describes the design and optimization of one configuration in this vehicle class, a lifting body with turbine-based low-speed propulsion. The integration of propulsion and airframe, both from an aero-propulsive and mechanical perspective are addressed. This paper primarily focuses on the design details of the preferred configuration and the analyses performed to assess its performance. The integration of both low-speed and high-speed propulsion is covered. Structural and mechanical designs are described along with materials and technologies used. Propellant and systems packaging are shown and the mission-sized vehicle weights are disclosed.

  17. Enhanced capability of the Combustion-Heated Scramjet Test Facility

    NASA Technical Reports Server (NTRS)

    Rock, Kenneth E.; Andrews, Earl H.; Eggers, James M.

    1991-01-01

    The Combustion-Heated Scramjet Test Facility (CHSTF) is described together with its modifications. The expanded simulation capabilities of the facility are documented. Nozzle exit surveys and tunnel calibration information are presented. It is noted that these modifications included a new heat-sink nickel liner heater, a new Mach 4.7 nozzle, and a new 70-ft vacuum sphere exhaust system. It is found that the facility in the air ejector mode of operation performed similarly to that prior to the addition of the vacuum sphere ducting.

  18. Expanded operational capabilities of the Langley Mach 7 Scramjet test facility

    NASA Technical Reports Server (NTRS)

    Thomas, S. R.; Guy, R. W.

    1983-01-01

    An experimental research program conducted to expand the operational capabilities of the NASA Langley Mach 7 Scramjet Test Facility is described. Previous scramjet testing in this facility was limited to a single simulated flight condition of Mach 6.9 at an altitude of 115,300 ft. The arc heater research demonstrates the potential of the facility for scramjet testing at simulated flight conditions from Mach 4 (at altitudes from 77,000 to 114,000 ft) to Mach 7 (at latitudes from 108,000 to 149,000 ft). Arc heater electrical characteristics, operational problems, measurements of nitrogen oxide contaminants, and total-temperature profiles are discussed.

  19. Massively parallel computation of three-dimensional scramjet combustor

    NASA Astrophysics Data System (ADS)

    Zheng, Z. H.; Le, J. L.

    Recent progress of computational study of scramjet combustor has been described in Refs 1-3. However, detailed flow properties, especially the lateral properties and the sidewall effects are not considered. In this paper, a parallel simulation of an experimental dual-mode scramjet combustor configuration is presented, considering the jet-to-jet symmetry and the full-duct modeling. Turbulence is modeled with the k-ɛ two-equation turbulence model and a 7-specie, 8-equation kinetics model is used to model hydrogen/air combustion. The conservation form of the Navier-Stokes equations with finite-rate chemistry reactions is solved using a diagonal implicit finite-volume method. For the two cases, the three-dimension flow-fields with equivalence ratio Φ=0.0 and 0.35 have been respectively simulated on the COW and MPP. Wall pressure comparisons between CFD and experiments (CARDC and NAL) show fair agreement for the jet-to-jet case. For the full-duct modeling, more detailed flow properties are obtained. The fuelpenetrating heights of the injectors are different because of the effects of the sidewall boundary layer and the shock wave in the combustor. According to numerical results, if adjusting the locations of the injectors, the combustion efficiency could be improved.

  20. Pylon effects on a scramjet cavity flameholder flowfield

    NASA Astrophysics Data System (ADS)

    Freeborn, Andrew B.

    Cavity flameholders in supersonic combustion ramjet (scramjet) combustors, while effective, fail to take advantage of the full combustor volume. Adding a pylon to the leading edge of a cavity flameholder generates a flowfield increasing mass exchange between the cavity and main combustor flow, increasing the mixing interface between flameholder products and main combustor flow, and exhibiting minimal Reynolds number effects. To demonstrate this modified flowfield driven by supersonic expansion behind the pylon, pylon-cavity flameholder flowfield data were obtained through a combination of wind tunnel experimentation and steady-state computational fluid dynamics (CFD). Flowfield effects of the pylon-cavity were examined at a Mach number of two and Reynolds numbers from approximately 32 million m-1 to 55 million m -1. Addition of the pylon resulted in approximately three times the mass exchange between the cavity and overlying flow. Both CFD and particle image velocimetry data showed strong upward flow behind the pylon, increasing exposure and exchange of cavity fluid with the main combustor flow. Reynolds number effects were weak within the range of test conditions. Assuming a suitably reacting fuel-air mixture, the addition of a pylon offers the scramjet designer an attractive option to take advantage of a greater proportion of combustor volume and improve combustor performance.

  1. MHD Energy Bypass Scramjet Performance with Real Gas Effects

    NASA Technical Reports Server (NTRS)

    Park, Chul; Mehta, Unmeel B.; Bogdanoff, David W.

    2000-01-01

    The theoretical performance of a scramjet propulsion system incorporating an magneto-hydro-dynamic (MHD) energy bypass scheme is calculated. The one-dimensional analysis developed earlier, in which the theoretical performance is calculated neglecting skin friction and using a sudden-freezing approximation for the nozzle flow, is modified to incorporate the method of Van Driest for turbulent skin friction and a finite-rate chemistry calculation in the nozzle. Unlike in the earlier design, in which four ramp compressions occurred in the pitch plane, in the present design the first two ramp compressions occur in the pitch plane and the next two compressions occur in the yaw plane. The results for the simplified design of a spaceliner show that (1) the present design produces higher specific impulses than the earlier design, (2) skin friction substantially reduces thrust and specific impulse, and (3) the specific impulse of the MHD-bypass system is still better than the non-MHD system and typical rocket over a narrow region of flight speeds and design parameters. Results suggest that the energy management with MHD principles offers the possibility of improving the performance of the scramjet. The technical issues needing further studies are identified.

  2. Modeling Shock Train Leading Edge Detection in Dual-Mode Scramjets

    NASA Astrophysics Data System (ADS)

    Ladeinde, Foluso; Lou, Zhipeng; Li, Wenhai

    2016-11-01

    The objective of this study is to accurately model the detection of shock train leading edge (STLE) in dual-mode scramjet (DMSJ) engines intended for hypersonic flight in air-breathing propulsion systems. The associated vehicles have applications in military warfare and intelligence, and there is commercial interest as well. Shock trains are of interest because they play a significant role in the inability of a DMSJ engine to develop the required propulsive force. The experimental approach to STLE detection has received some attention; as have numerical calculations. However, virtually all of the numerical work focus on mechanically- (i.e., pressure-) generated shock trains, which are much easier to model relative to the phenomenon in the real system where the shock trains are generated by combustion. A focus on combustion, as in the present studies, enables the investigation of the effects of equivalence ratio, which, together with the Mach number, constitutes an important parameter determining mode transition. The various numerical approaches implemented in our work will be reported, with result comparisons to experimental data. The development of an STLE detection procedure in an a priori manner will also be discussed.

  3. Design Considerations of ISTAR Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

    NASA Technical Reports Server (NTRS)

    Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

    2003-01-01

    The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system that produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

  4. Design Considerations of Istar Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

    NASA Technical Reports Server (NTRS)

    Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

    2002-01-01

    The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system thai: produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

  5. Supersonic combustion engine testbed, heat lightning

    NASA Technical Reports Server (NTRS)

    Hoying, D.; Kelble, C.; Langenbahn, A.; Stahl, M.; Tincher, M.; Walsh, M.; Wisler, S.

    1990-01-01

    The design of a supersonic combustion engine testbed (SCET) aircraft is presented. The hypersonic waverider will utilize both supersonic combustion ramjet (SCRAMjet) and turbofan-ramjet engines. The waverider concept, system integration, electrical power, weight analysis, cockpit, landing skids, and configuration modeling are addressed in the configuration considerations. The subsonic, supersonic and hypersonic aerodynamics are presented along with the aerodynamic stability and landing analysis of the aircraft. The propulsion design considerations include: engine selection, turbofan ramjet inlets, SCRAMjet inlets and the SCRAMjet diffuser. The cooling requirements and system are covered along with the topics of materials and the hydrogen fuel tanks and insulation system. A cost analysis is presented and the appendices include: information about the subsonic wind tunnel test, shock expansion calculations, and an aerodynamic heat flux program.

  6. The Effect of Pulsed Injection on Shear Layer Dynamics in a Scramjet Combustion Chamber

    NASA Astrophysics Data System (ADS)

    Smith, Leslie

    One of the greatest problems that scramjet research faces is fuel air mixing. The residence time for a scramjet engine, or the time it takes for a volume of air to completely pass through the engine, is on the order of 0.1 ms. In that extremely short period of time fuel must be injected and fully mirco-mixed at stoichiometric ratios with the combustion chamber airflow. The fuel-air mixture must then be combusted and expanded through the nozzle to produce thrust. The goal of this research is to develop a new more efficient method of fuel air mixing within a scramjet combustion chamber. A possible way to speed up the mixing process of parallel injection without incurring the total pressure losses that would occur in normal injection is to inject the fuel from the rear side of a backward facing step. Backward facing steps in supersonic flow produce a Prandtl-Meyer expansion fan followed by a shear layer. The instabilities in this shear layer have dominant resonant frequencies. It is believed that if fuel is injected in pulses that impinge on the shear layer at these dominant resonant frequencies that the shear layer will resonate. When the shear layer resonates the vortices that form in the shear layer will grow in magnitude, thus mixing the injected fuel with the air. To test this hypothesis a new test section was designed and built that features a one inch step under which an injector can be housed. This new test section was installed in the supersonic facility at the University of Kansas. Two injectors were also designed that each feature a face plate, one with eight injection ports arranged in a ring and one with 5 injection ports. Between the face plate and a back plate there is a cavity that houses a rotating valve that is powered by a pneumatic motor. Five valves were built: one with 8 teeth, one with 16 teeth, one with 5 teeth that are the same size as the gaps between the teeth, one with 5 teeth where the teeth are 50% larger than the gaps, and one with 5

  7. Review of HxPyOz-Catalyzed H + OH Recombination in Scramjet Nozzle Expansions; and Possible Phosphoric Acid Enhancement of Scramjet Flameholding, from Extinction of H3PO4 + H2 - Air Counterflow Diffusion Flames

    NASA Technical Reports Server (NTRS)

    Pellett, Gerald

    2005-01-01

    Recent detailed articles by Twarowski indicate that small quantities of phosphorus oxides and acids in the fuel-rich combustion products of H2 + phosphine (PH3) + air should significantly catalyze H, OH and O recombination kinetics during high-speed nozzle expansions -- to reform H2O, release heat, and approach equilibrium more rapidly and closely than uncatalyzed kinetics. This paper is an initial feasibility study to determine (a) if addition of phosphoric acid vapor (H3PO4) to a H2 fuel jet -- which is much safer than using PH3 -- will allow combustion in a high-speed scramjet engine test without adverse effects on localized flameholding, and (b) if phosphorus-containing exhaust emissions are environmentally acceptable. A well-characterized axisymmetric straight-tube opposed jet burner (OJB) tool is used to evaluate H3PO4 addition effects on the air velocity extinction limit (flame strength) of a H2 versus air counterflow diffusion flame. Addition of nitric oxide (NO), also believed to promote catalytic H-atom recombination, was evaluated for comparison. Two to five mass percent H3PO4 in the H2 jet increased flame strength 4.2%, whereas airside addition decreased it 1%. Adding 5% NO to the H2 caused a 2% decrease. Products of H-atom attack on H3PO4 produced an intense green chemiluminescence near the stagnation point. The resultant exothermic production of phosphorus oxides and acids, with accelerated H-atom recombination, released sufficient heat near the stagnation point to increase flame strength. In conclusion, the addition of H3PO4 vapor (or more reactive P sources) to hydrogen in scramjet engine tests may positively affect flameholding stability in the combustor and thrust production during supersonic expansion -- a possible dual benefit with system design / performance implications. Finally, a preliminary assessment of possible environmental effects indicates that scramjet exhaust emissions should consist of phosphoric acid aerosol, with gradual

  8. Development of a Test to Evaluate Aerothermal Response of Materials to Hypersonic Flow Using a Scramjet Wind Tunnel (Postprint)

    DTIC Science & Technology

    2010-05-01

    materials under hypersonic flow conditions was developed and evaluated. A direct-connect scramjet combustor rig, de- signed based on the needs for...hypersonic flight and de- scribed elsewhere,10 was explored for use as a wind tunnel. A methodology to introduce samples into the flow path of the combustor ...Scramjet Direct-Connect Rig Gruber et al.10 have designed and fabricated a di- rect-connect full-scale scramjet combustor test facility for studies on

  9. Status of the Combined Cycle Engine Rig

    NASA Technical Reports Server (NTRS)

    Saunders, Dave; Slater, John; Dippold, Vance

    2009-01-01

    Status for the past year is provided of the turbine-based Combined-Cycle Engine (CCE) Rig for the hypersonic project. As part of the first stage propulsion of a two-stage-to-orbit vehicle concept, this engine rig is designed with a common inlet that supplies flow to a turbine engine and a dual-mode ramjet / scramjet engine in an over/under configuration. At Mach 4 the inlet has variable geometry to switch the airflow from the turbine to the ramjet / scramjet engine. This process is known as inlet mode-transition. In addition to investigating inlet aspects of mode transition, the rig will allow testing of turbine and scramjet systems later in the test series. Fully closing the splitter cowl "cocoons" the turbine engine and increases airflow to the scramjet duct. The CCE Rig will be a testbed to investigate integrated propulsion system and controls technology objectives. Four phases of testing are planned to 1) characterize the dual inlet database, 2) collect inlet dynamics using system identification techniques, 3) implement an inlet control to demonstrate mode-transition scenarios and 4) demonstrate integrated inlet/turbine engine operation through mode-transition. Status of the test planning and preparation activities is summarized with background on the inlet design and small-scale testing, analytical CFD predictions and some details of the large-scale hardware. The final stages of fabrication are underway.

  10. Numerical Simulation of Dual-Mode Scramjet Combustors

    NASA Technical Reports Server (NTRS)

    Rodriguez, C. G.; Riggins, D. W.; Bittner, R. D.

    2000-01-01

    Results of a numerical investigation of a three-dimensional dual-mode scramjet isolator-combustor flow-field are presented. Specifically, the effect of wall cooling on upstream interaction and flow-structure is examined for a case assuming jet-to-jet symmetry within the combustor. Comparisons are made with available experimental wall pressures. The full half-duct for the isolator-combustor is then modeled in order to study the influence of side-walls. Large scale three-dimensionality is observed in the flow with massive separation forward on the side-walls of the duct. A brief review of convergence-acceleration techniques useful in dual-mode simulations is presented, followed by recommendations regarding the development of a reliable and unambiguous experimental data base for guiding CFD code assessments in this area.

  11. Flow enthalpy effects on scramjet mixing and combustion

    NASA Technical Reports Server (NTRS)

    Riggins, D. W.; Mcclinton, C. R.; Rogers, R. C.

    1992-01-01

    The present analysis of the 3D characteristics of the mixing processes in a scramjet combustor gives attention to the streamwise, vorticity-driven macromixing, as well as shear-driven small-scale mixing, in both an unconfined Mach 6 airstream and a high-enthalpy (Mach-17) confined, Mach 6 airstream. The results obtained indicate that mixing is substantially lower for the high-enthalpy case. The effect of residence time on jet mixing is discussed in view of the production and decay of the axial vorticity, cross-flow velocities, and mean-flow velocities of these confined flows. The effective distance travelled by a fluid particle in the cross-flow, over the combustor length, is used to explain axial vorticity's contribution to mixing.

  12. Analysis of a hydrocarbon scramjet with augmented preburning

    NASA Technical Reports Server (NTRS)

    Molvik, Gregory A.; Bowles, Jeffrey V.; Huynh, Loc C.

    1992-01-01

    This paper presents the results of a feasibility study of a hydrocarbon scramjet design utilizing an augmented preburner upstream of the main fuel injector locations. The combustor design evaluated here is for a small hypersonic research vehicle. It consists of a preburner into which a small amount of fuel is burned with on-board liquid oxygen and injected into the main airflow, upstream of the main fuel injector locations, thus ensuring that combustion is present and uninterrupted. Two degrees of analysis are presented including a one-dimensional cycle analysis and a complete computational fluid dynamic analysis with finite-rate chemistry and a two-equation turbulence model. Comparison of these analyses show good agreement when the CFD-predicted fuel consumption schedule is used in the cycle analysis.

  13. O and temperature in a hydrocarbon-fueled scramjet combustor

    NASA Astrophysics Data System (ADS)

    Goldenstein, C. S.; Schultz, I. A.; Spearrin, R. M.; Jeffries, J. B.; Hanson, R. K.

    2014-09-01

    The design and demonstration of a two-color tunable diode laser sensor for measurements of temperature and H2O in an ethylene-fueled model scramjet combustor are presented. This sensor probes multiple H2O transitions in the fundamental vibration bands near 2.5 μm that are up to 20 times stronger than those used by previous near-infrared H2O sensors. In addition, two design measures enabled high-fidelity measurements in the nonuniform flow field. (1) A recently developed calibration-free scanned-wavelength-modulation spectroscopy spectral-fitting strategy was used to infer the integrated absorbance of each transition without a priori knowledge of the absorption lineshape and (2) transitions with strengths that scale near-linearly with temperature were used to accurately determine the H2O column density and the H2O-weighted path-averaged temperature from the integrated absorbance of two transitions.

  14. Investigation of the feasibility of CARS measurements in scramjet combustion

    NASA Technical Reports Server (NTRS)

    Shirley, J. A.; Hall, R. J.; Eckbreth, A. C.

    1980-01-01

    Results are presented of analytical and experimental investigations to determine the feasibility of using coherent anti-Stokes Raman Spectroscopy (CARS) to measure temperature and species concentration in supersonic combustion experiments. The CARS spectra of H2O, O2 and H2 were measured in laboratory flames. Computer code calculated spectra agree very well with the measured spectra. Temperature, and O2 and H2 concentration profiles have been determined from CARS spectra in a laboratory H2 air flat diffusion flame. Temperature measurements agree with radiation corrected thermocouple measurements within 5 to 10 percent, depending on species concentration. The feasibility of measuring O2 concentrations up to 10 percent, from the spectral shape was demonstrated. H2 concentrations determined from CARS intensities agree with spontaneous Raman measurements within a factor of two. Finally, a conceptual design was formulated for diagnostics in the Langley Research Center scramjet combustion facility.

  15. A Priori Analysis of Flamelet-Based Modeling for a Dual-Mode Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Quinlan, Jesse R.; McDaniel, James C.; Drozda, Tomasz G.; Lacaze, Guilhem; Oefelein, Joseph

    2014-01-01

    An a priori investigation of the applicability of flamelet-based combustion models to dual-mode scramjet combustion was performed utilizing Reynolds-averaged simulations (RAS). For this purpose, the HIFiRE Direct Connect Rig (HDCR) flowpath, fueled with a JP-7 fuel surrogate and operating in dual- and scram-mode was considered. The chemistry of the JP-7 fuel surrogate was modeled using a 22 species, 18-step chemical reaction mechanism. Simulation results were compared to experimentally-obtained, time-averaged, wall pressure measurements to validate the RAS solutions. The analysis of the dual-mode operation of this flowpath showed regions of predominately non-premixed, high-Damkohler number, combustion. Regions of premixed combustion were also present but associated with only a small fraction of the total heat-release in the flow. This is in contrast to the scram-mode operation, where a comparable amount of heat is released from non-premixed and premixed combustion modes. Representative flamelet boundary conditions were estimated by analyzing probability density functions for temperature and pressure for pure fuel and oxidizer conditions. The results of the present study reveal the potential for a flamelet model to accurately model the combustion processes in the HDCR and likely other high-speed flowpaths of engineering interest.

  16. Shock Train/Boundary-Layer Interaction in Rectangular Scramjet Isolators

    NASA Astrophysics Data System (ADS)

    Geerts, Jonathan Simon

    Numerous studies of the dual-mode scramjet isolator, a critical component in preventing inlet unstart and/or vehicle loss by containing a collection of flow disturbances called a shock train, have been performed since the dual-mode propulsion cycle was introduced in the 1960s. Low momentum corner flow and other three-dimensional effects inherent to rectangular isolators have, however, been largely ignored in experimental studies of the boundary layer separation driven isolator shock train dynamics. Furthermore, the use of two dimensional diagnostic techniques in past works, be it single-perspective line-of-sight schlieren/shadowgraphy or single axis wall pressure measurements, have been unable to resolve the three-dimensional flow features inside the rectangular isolator. These flow characteristics need to be thoroughly understood if robust dual-mode scramjet designs are to be fielded. The work presented in this thesis is focused on experimentally analyzing shock train/boundary layer interactions from multiple perspectives in aspect ratio 1.0, 3.0, and 6.0 rectangular isolators with inflow Mach numbers ranging from 2.4 to 2.7. Secondary steady-state Computational Fluid Dynamics studies are performed to compare to the experimental results and to provide additional perspectives of the flow field. Specific issues that remain unresolved after decades of isolator shock train studies that are addressed in this work include the three-dimensional formation of the isolator shock train front, the spatial and temporal low momentum corner flow separation scales, the transient behavior of shock train/boundary layer interaction at specific coordinates along the isolator's lateral axis, and effects of the rectangular geometry on semi-empirical relations for shock train length prediction. (Abstract shortened by ProQuest.).

  17. Performance of a RBCC Engine in Rocket-Operation

    NASA Astrophysics Data System (ADS)

    Tomioka, Sadatake; Kubo, Takahiro; Noboru Sakuranaka; Tani, Koichiro

    Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket-based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

  18. The effect of exhaust plume/afterbody interaction on installed Scramjet performance

    NASA Technical Reports Server (NTRS)

    Edwards, Thomas Alan

    1988-01-01

    Newly emerging aerospace technology points to the feasibility of sustained hypersonic flight. Designing a propulsion system capable of generating the necessary thrust is now the major obstacle. First-generation vehicles will be driven by air-breathing scramjet (supersonic combustion ramjet) engines. Because of engine size limitations, the exhaust gas leaving the nozzle will be highly underexpanded. Consequently, a significant amount of thrust and lift can be extracted by allowing the exhaust gases to expand along the underbody of the vehicle. Predicting how these forces influence overall vehicle thrust, lift, and moment is essential to a successful design. This work represents an important first step toward that objective. The UWIN code, an upwind, implicit Navier-Stokes computer program, has been applied to hypersonic exhaust plume/afterbody flow fields. The capability to solve entire vehicle geometries at hypersonic speeds, including an interacting exhaust plume, has been demonstrated for the first time. Comparison of the numerical results with available experimental data shows good agreement in all cases investigated. For moderately underexpanded jets, afterbody forces were found to vary linearly with the nozzle exit pressure, and increasing the exit pressure produced additional nose-down pitching moment. Coupling a species continuity equation to the UWIN code enabled calculations indicating that exhaust gases with low isentropic exponents (gamma) contribute larger afterbody forces than high-gamma exhaust gases. Moderately underexpanded jets, which remain attached to unswept afterbodies, underwent streamwise separation on upswept afterbodies. Highly underexpanded jets produced altogether different flow patterns, however. The highly underexpanded jet creates a strong plume shock, and the interaction of this shock with the afterbody was found to produce complicated patterns of crossflow separation. Finally, the effect of thrust vectoring on vehicle balance has

  19. Method for obtaining aerodynamic data on hypersonic configurations with scramjet exhaust flow simulation

    NASA Technical Reports Server (NTRS)

    Hartill, W. R.

    1977-01-01

    A hypersonic wind tunnel test method for obtaining credible aerodynamic data on a complete hypersonic vehicle (generic X-24c) with scramjet exhaust flow simulation is described. The general problems of simulating the scramjet exhaust as well as accounting for scramjet inlet flow and vehicle forces are analyzed, and candidate test methods are described and compared. The method selected as most useful makes use of a thrust-minus-drag flow-through balance with a completely metric model. Inlet flow is diverted by a fairing. The incremental effect of the fairing is determined in the testing of two reference models. The net thrust of the scramjet module is an input to be determined in large-scale module tests with scramjet combustion. Force accounting is described, and examples of force component levels are predicted. Compatibility of the test method with candidate wind tunnel facilities is described, and a preliminary model mechanical arrangement drawing is presented. The balance design and performance requirements are described in a detailed specification. Calibration procedures, model instrumentation, and a test plan for the model are outlined.

  20. Parametric study of combustion oscillation in a single-side expansion scramjet combustor

    NASA Astrophysics Data System (ADS)

    Ouyang, Hao; Liu, Weidong; Sun, Mingbo

    2016-10-01

    As a promising candidate for future air-breathing systems, the viability and efficiency of scramjet propulsion is challenged by a variety of factors including the combustion oscillation in scramjet combustor. A series of comparative experiments focusing on the combustion oscillation issue has been carried out in the present work. The obtained experimental results show that as the global equivalence ratio increases, the combustion oscillation becomes more regular and frequent which is the most intensive in the vicinity of the fuel jet and the periodic combustion oscillation is more possible when the injectors and flame-holding cavity are mounted on the expansion-side wall. In order to avoid the combustion oscillation in scramjet combustor, distributed injection scheme is an effective method which can induce two parts interacting stable flame. In addition, the results reveal that the varying fuel including hydrogen, ethylene and kerosene with different chemical kinetics has a significant effect on the reaction process in scramjet combustor, which can result in stable combustion, periodic oscillation and failed ignition respectively on the same operating condition of this paper. We believe that the present work is helpful to the designing of scramjet propulsion device.

  1. Thermal performance of a scramjet combustor operating at Mach 5.6 flight conditions. Final report, May 1996--May 1997

    SciTech Connect

    Stouffer, S.D.; Neumann, R.D.; Emmer, D.S.

    1997-10-01

    This report describes the experimental data and the procedures used in acquiring and reducing the thermal loads data during tests of a hydrocarbon-fueled scramjet combustor at United Technologies Research Center (UTRC). This research effort is part of the UTRC effort to develop dual-mode scramjet combustor technology to support the development of Mach S missile technology. The objective of the thermal loads testing was to map the thermal and mechanical loads, including heat transfer, dynamic and static pressures, and skin friction in a scramjet combustor during direct-connect scramjet tests. The tests were conducted at the UTRC Ramject/Scramjet direct-connect combustor test facility in East Hartford, CT.

  2. Autonomous space processor for orbital debris removal and flame augmentation additives in scramjets for the National Aerospace Plane

    NASA Technical Reports Server (NTRS)

    1988-01-01

    This is a brief description of the USRA-sponsored design project at the University of Arizona. Approximately eighty-percent of this effort was spent pursuing a novel engineering concept for the in-situ processing of orbital debris utilizing resources available in low Earth orbit (LEO); the other twenty-percent was devoted to discovering innovative additives for the anchoring of supersonic combustion zones that find direct use in the Aerospace Plane that is expected to use scramjets. The seriousness of the orbital debris problem is briefly described. Available 'solutions' are outlined from the literature. The engineering design is briefly mentioned, with an emphasis on the positive aspects of the space environment that should be used in an economical approach. The aspects of operating in microgravity, vacuum, and in utilizing solar energy are mentioned. A quantitative computer animation was developed to provide design data. Three specific dead spacecraft were identified for an initial cleanup mission. The design concept, which includes a solar processor, remote arm manipulators, and the gradual processing of the debris, is also described. This is followed by a description of hardware construction. Operation and actual processing of simulated debris parts (aluminum, for now) are demonstrated in the NASP task, construction of the new design for measuring the radiation from the key free radicals (as enhanced by the additives) is described. Immediate (1988) and long-range (through 1992) future plans are shown to clearly indicate the full engineering design strategy in the light of the national space program thrusts.

  3. Review of the potential of silanes as rocket/scramjet fuels

    NASA Astrophysics Data System (ADS)

    Hidding, Bernhard; Pfitzner, Michael; Simone, Domenico; Bruno, Claudio

    2008-07-01

    Experimental use as well as theoretical considerations regarding silanes as fuels for spacecrafts and supersonic flight are reviewed. The historical circumstances leading to the utilization of monosilane as a fuel additive for scramjets are highlighted and milestones in the chemical research on silanes are summarized. Recent developments such as the use of monosilane as an ignition aid in the NASA X-43A scramjet flights as well as general progress in silicon hydride research, including liquid higher silanes and the resulting potential for the propulsion field are discussed.

  4. Experimental and Computational Study of Flow Interactions in a Generic Scramjet Inlet-Isolator

    NASA Astrophysics Data System (ADS)

    Idris, A. Che; Saad, M. R.; Zare-Behtash, H.; Kontis, K.

    A supersonic combustion ramjet (scramjet) inlet that is optimized for cruise conditions would experience an adverse flow conditions at off-design conditions. Strong viscous interactions, shock-shock interactions, glancing shock interactions, shockboundary layer interactions, flow separations and other problems could occur during such conditions [1][2].

  5. A Two-Line Absorption Instrument for Scramjet Temperature and Water Vapor Concentration Measurement in HYPULSE

    NASA Technical Reports Server (NTRS)

    Tsai, C. Y.

    1998-01-01

    A three beam water vapor sensor system has been modified to provide for near simultaneous temperature measurement. The system employs a tunable diode laser to scan spectral line of water vapor. The application to measurements in a scramjet combustor environment of a shock tunnel facility is discussed. This report presents and discusses die initial calibration of the measurement system.

  6. Airframe Integration Trade Studies for a Reusable Launch Vehicle

    NASA Technical Reports Server (NTRS)

    Dorsey, John T.; Wu, Chauncey; Rivers, Kevin; Martin, Carl; Smith, Russell

    1999-01-01

    Future launch vehicles must be lightweight, fully reusable and easily maintained if low-cost access to space is to be achieved. The goal of achieving an economically viable Single-Stage-to-Orbit (SSTO) Reusable Launch Vehicle (RLV) is not easily achieved and success will depend to a large extent on having an integrated and optimized total system. A series of trade studies were performed to meet three objectives. First, to provide structural weights and parametric weight equations as inputs to configuration-level trade studies. Second, to identify, assess and quantify major weight drivers for the RLV airframe. Third, using information on major weight drivers, and considering the RLV as an integrated thermal structure (composed of thrust structures, tanks, thermal protection system, insulation and control surfaces), identify and assess new and innovative approaches or concepts that have the potential for either reducing airframe weight, improving operability, and/or reducing cost.

  7. Propulsion-airframe integration for commercial and military aircraft

    NASA Technical Reports Server (NTRS)

    Henderson, William P.

    1988-01-01

    A significant level of research is ongoing at NASA's Langley Research Center on integrating the propulsion system with the aircraft. This program has included nacelle/pylon/wing integration for turbofan transports, propeller/nacelle/wing integration for turboprop transports, and nozzle/afterbody/empennage integration for high performance aircraft. The studies included in this paper focus more specifically on pylon shaping and nacelle location studies for turbofan transports, nacelle and wing contouring and propeller location effects for turboprop transports, and nozzle shaping and empennage effects for high performance aircraft. The studies were primarily conducted in NASA Langley's 16-Foot Transonic Tunnel at Mach numbers up to 1.20. Some higher Mach number data obtained at NASA's Lewis Research Center is also included.

  8. Engine Validation of Noise and Emission Reduction Technology Phase I

    NASA Technical Reports Server (NTRS)

    Weir, Don (Editor)

    2008-01-01

    This final report has been prepared by Honeywell Aerospace, Phoenix, Arizona, a unit of Honeywell International, Inc., documenting work performed during the period December 2004 through August 2007 for the NASA Glenn Research Center, Cleveland, Ohio, under the Revolutionary Aero-Space Engine Research (RASER) Program, Contract No. NAS3-01136, Task Order 8, Engine Validation of Noise and Emission Reduction Technology Phase I. The NASA Task Manager was Dr. Joe Grady of the NASA Glenn Research Center. The NASA Contract Officer was Mr. Albert Spence of the NASA Glenn Research Center. This report is for a test program in which NASA funded engine validations of integrated technologies that reduce aircraft engine noise. These technologies address the reduction of engine fan and jet noise, and noise associated with propulsion/airframe integration. The results of these tests will be used by NASA to identify the engineering tradeoffs associated with the technologies that are needed to enable advanced engine systems to meet stringent goals for the reduction of noise. The objectives of this program are to (1) conduct system engineering and integration efforts to define the engine test-bed configuration; (2) develop selected noise reduction technologies to a technical maturity sufficient to enable engine testing and validation of those technologies in the FY06-07 time frame; (3) conduct engine tests designed to gain insight into the sources, mechanisms and characteristics of noise in the engines; and (4) establish baseline engine noise measurements for subsequent use in the evaluation of noise reduction.

  9. Investigations on the Influence of the In-Stream Pylon and Strut on the Performance of a Scramjet Combustor

    PubMed Central

    Liu, Weidong; Sun, Mingbo

    2014-01-01

    The influence of the in-stream pylon and strut on the performance of scramjet combustor was experimentally and numerically investigated. The experiments were conducted with a direct-connect supersonic model combustor equipped with multiple cavities. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The research results show that, compared with the scramjet combustor without pylon and strut, the wall pressure and the thrust of the scramjet increase due to the improvement of mixing and combustion effect due to the pylon and strut. The total pressure loss caused by the strut is considerable whereas pylon influence is slight. PMID:25254234

  10. Experimental/Computational Studies of Combined Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors

    DTIC Science & Technology

    2010-10-01

    Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors 5a. CONTRACT NUMBER 5b. GRANT NUMBER FA9550-04-1-0387 5c...development of new tunable diode-laser diagnostics for supersonic combustors . The scientific results obtained from this MURI program were disseminated widely...Inlets and Scramjet Combustors TABLE OF CONTENTS TOPIC PAGE 1. COVER PAGE / ABSTRACT i 1. SUMMARY OF ACTIVITIES 1 2. PERSONNEL SUPPORTED 2 3

  11. Experimental/Computational Studies of Combined-Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors

    DTIC Science & Technology

    2010-05-22

    Scramjet Combustors 5b. GRANT NUMBER FA9550-04-1-0387 5c. PROGRAM ELEMENT NUMBER 61103F 6. AUTHOR(S) 5d. PROJECT NUMBER 5094...technique for compressible jets-in-crossflow, and development of new tunable diode-laser diagnostics for supersonic combustors . The scientific results...Computational Studies of Combined-Cycle Propulsion: Physics and Transient Phenomena in Inlets and Scramjet Combustors TABLE OF CONTENTS

  12. Survey of supersonic combustion ramjet research at Langley

    NASA Technical Reports Server (NTRS)

    Northam, G. B.; Anderson, G. Y.

    1986-01-01

    The Hypersonic Propulsion Branch at NASA Langley Research Center has maintained an active research program in supersonic combustion ramjet (scramjet) and high speed ramjet propulsion since the 1960s. The focus for this research has centered on propulsion for manned reuseable vehicles with cryogenic hydrogen fuel. This paper presents some highlights of this research. The design philosophy of the Langley fixed-geometry airframe-integrated modular scramjet is discussed. The component development and research programs that have supported the successful demonstration of the engine concept using subscale engine module hardware is reviewed and a brief summary of the engine tests presented. An extensive bibliography of research supported by the Langley program is also included.

  13. Comparison of Performances of Scramjet-Driven Experimental DCW-MHD Generators with Different Cross-Section

    NASA Astrophysics Data System (ADS)

    Niwa, Naoyuki; Takahashi, Toru; Fujino, Takayasu; Ishikawa, Motoo

    The purpose of this study is to examine the influence of shape of cross-section of scramjet engine driven experimental DCW-MHD generator on generator performance by three-dimensional numerical analyses. We have designed the MHD generators with symmetric square and circular cross-section, based on the experimental MHD generator with asymmetric square cross-section. Under the optimum load condition, the electric power output becomes 26.6kW for the asymmetric square cross-section, 24.6kW for the symmetric square cross-section, and 22.4kW for the circular cross-section. The highest output is obtained for the experimental generator with asymmetric square cross-section. The difference of electric power output is induced by the difference of flow velocity and boundary layer thickness. For the generator with asymmetric square cross-section, the average flow velocity becomes the highest and the boundary layer becomes the thinnest. The compression wave is generated depending on the channel shape. The difference of flow velocity and boundary layer thickness is induced by the superposition of compression wave.

  14. Exit plane H2O concentration measurements correlated with OH PLIF near-injector mixing measurements for scramjet flows

    NASA Technical Reports Server (NTRS)

    Parker, T. E.; Allen, Mark G.; Foutter, R. R.; Sonnenfroh, D. M.; Rawlins, W. T.

    1992-01-01

    Mixing and combusting high enthalpy flows, similar to those encountered in scramjet engines, were investigated using a shock tunnel to produce the flow in conjunction with non-intrusive optical diagnostics which monitored the performance of two injector configurations. The shock tunnel is configured to produce Mach 3 flow and stagnation enthalpies corresponding to flight equivalent Mach numbers between 7 and 11. A pulsed hydrogen injection capability and interchangeable injector blocks provide a means of examining high speed, high enthalpy reacting flows. Planar laser induced fluorescence (PLIF) of OH molecules in the near injector region produced images which show the combusting and mixing zones for the reacting flow. Line-of-sight exit plane measurement of water concentration and temperature were used to provide a unique method of monitoring exit plane products. These results demonstrated that a velocity matched axial injection system produced a fuel jet that lifted off the floor of the duct. Mixing was observed to increase for this system as a velocity mismatch was introduced. Comparison of exit plane water concentrations for a wall jet injection system and a velocity matched injection system indicated similar mixing performance but an accurate pressure measurement is necessary to further validate the result. In addition, exit plane measurements indicated an approximate steady-state condition was achieved during the 1 to 2 ms test times.

  15. Scramjet Combustor Simulations Using Reduced Chemical Kinetics for Practical Fuels

    DTIC Science & Technology

    2003-12-01

    Stochastic Simulation of an HCCI Engine Using an Automatically Reduced Mechanism,� ICE, Vol. 37-2, 2001 Fall Technical Conference...Christopher J. Montgomery, and Wei Zhao Reaction Engineering International (REI) 77 West 200 South, Suite 210 Salt Lake City, UT 84101 Dean R...DOUGLAS L. DAVIS CAPT. BRIAN C. MCDONALD AFRL/PRAS, Project Engineer Branch Chief Propulsion Sciences Branch Propulsion

  16. Experimental and computational study of the effect of shocks on film cooling effectiveness in scramjet combustors

    NASA Technical Reports Server (NTRS)

    Kamath, Pradeep S.; Holden, Michael S.; Mcclinton, Charles R.

    1990-01-01

    This paper presents results from a study conducted to investigate the effect of incident oblique shocks on the effectiveness of a coolant film at Mach numbers, typical of those expected in a scramjet combustor at Mach 15 to 20 flight. Computations with a parabolic code are in good agreement with the measured pressures and heat fluxes, after accounting for the influence of the shock upstream of its point of impingement on the plate, and the expansion from the trailing edge of the shock generator. The test data shows that, for the blowing rates tested, the film is rendered largely ineffective by the shock. Computations show that coolant blowing rates five to ten times those tested are required to protect against shock-induced heating. The implications of the results to scramjet combustor design are discussed.

  17. Simulation of hypersonic scramjet exhaust. [pressure distribution on afterbody/nozzle sections of vehicle

    NASA Technical Reports Server (NTRS)

    Oman, R. A.; Foreman, K. M.; Leng, J.; Hopkins, H. B.

    1975-01-01

    A plan and some preliminary analysis for the accurate simulation of pressure distributions on the afterbody/nozzle portions of a hypersonic scramjet vehicle are described. The objectives fulfilled were to establish the standards of similitude for a hydrogen/air scramjet exhaust interacting with a vehicle afterbody, determine an experimental technique for validation of the procedures that will be used in conventional wind tunnel facilities, suggest a program of experiments for proof of the concept, and explore any unresolved problems in the proposed simulation procedures. It is shown that true enthalpy, Reynolds number, and nearly exact chemistry can be provided in the exhaust flow for the flight regime from Mach 4 to 10 by a detonation tube simulation. A detailed discussion of the required similarity parameters leads to the conclusion that substitute gases can be used as the simulated exhaust gas in a wind tunnel to achieve the correct interaction forces and moments.

  18. Nonlinear characteristics and detection of combustion modes for a hydrocarbon fueled scramjet

    NASA Astrophysics Data System (ADS)

    Zhang, Cong; Yang, Qingchun; Chang, Juntao; Tang, Jingfeng; Bao, Wen

    2015-05-01

    An experimental investigation of combustion mode transition in a hydrocarbon fueled scramjet combustor model is reported under Mach number 2.1 and 2.5 inflow conditions. Three different combustion modes with respect to equivalence ratio are observed, namely, scramjet mode, weak ramjet mode and strong ramjet mode. The typical features of different combustion modes are analyzed by wall-pressures and one-dimensionally estimated Mach number distributions. The processes of combustion mode transitions show significant nonlinear characteristics. The static pressure and Mach number have discontinuous sudden changes as the mode transition occurs, especially near the fuel-supply region, emphasizing the importance of detection and control of combustion modes. The nonlinear characteristics of wall-pressures near the exit of the isolator can be used in the detection of different combustion modes. A series of experiments prove that this pressure-magnitude-based detection technique is feasible.

  19. Dual-Pump CARS Measurements in the University of Virginia's Dual-Mode Scramjet: Configuration "A"

    NASA Technical Reports Server (NTRS)

    Cutler, Andrew D.; Magnotti, Gaetano; Gallo, Emanuela; Danehy, Paul M.; Rockwell, Robert; Goyne, Christopher P.; McDaniel, James

    2012-01-01

    In this paper we describe efforts to obtain canonical data sets to assist computational modelers in their development of models for the prediction of mixing and combustion in scramjet combustors operating in the ramjet-scramjet transition regime. The CARS technique is employed to acquire temporally and spatially resolved measurements of temperature and species mole-fraction at four planes, one upstream of an H2 fuel injector and three downstream. The technique is described and results are presented for cases with and without chemical reaction. The vibrational energy mode in the heated airstream of the combustor was observed to be frozen at near facility heater conditions and significant nonuniformities in temperature were observed, attributed to nonuniformities of temperature exiting the heater. The measurements downstream of fuel injection show development of mixing and combustion, and are already proving useful to the modelers.

  20. Hyper-X and Pegasus Launch Vehicle: A Three-Foot Model of the Hypersonic Experimental Research Vehic

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The configuration of the X-43A Hypersonic Experimental Research Vehicle, or Hyper-X, attached to a Pegasus launch vehicle is displayed in this three-foot-long model at NASA's Dryden Flight Research Center, Edwards, California. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43

  1. X-43A Undergoing Controlled Radio Frequency Testing in the Benefield Anechoic Facility at Edwards Ai

    NASA Technical Reports Server (NTRS)

    2000-01-01

    The X-43A Hypersonic Experimental (Hyper-X) Vehicle hangs suspended in the cavernous Benefield Aenechoic Facility at Edwards Air Force Base during radio frequency tests in January 2000. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is the airframe integration

  2. Flowfield measurements in a model scramjet combustion using laser-induced iodine fluorescence

    NASA Technical Reports Server (NTRS)

    Mcdaniel, J. C., Jr.

    1984-01-01

    Preliminary designs were completed for an iodine mixing chamber and the optical setup to be used with a modified wind tunnel in obtaining accurate, spatially resolved measurements of variables in the flowfield of a model nonreacting scramjet combustor. Schematics of the iodine-seeded wind tunnel and a sketch of the charcoal filter for removing the iodine are included along with a cutaway section of the laboratory.

  3. A source flow characteristic technique for the analysis of scramjet exhaust flow field

    NASA Technical Reports Server (NTRS)

    Delguidice, P.; Dash, S.; Kalben, P.

    1974-01-01

    The factors which influence the design and selection of a nozzle for a hypersonic scramjet are described. A two dimensional second-order characteristic procedure capable of analyzing the aerodynamic performance of typical nozzle configurations is presented. Equations of motion governing the two dimensional, axisymmetric, or axially expanding inviscid flow of a gas mixture, with frozen chemistry, are provided. Diagrams of the flow conditions for various configurations are included.

  4. Nitric Oxide and Oxygen Air-Contamination Effects on Extinction Limits of Non-Premixed Hydrocarbon-Air Flames for a HIFiRE Scramjet

    NASA Technical Reports Server (NTRS)

    Pellett, Gerald L.; Dawson, Lucy C.; Vaden, Sarah N.; Wilson, Lloyd G.

    2009-01-01

    depletion (from 21 to 19.5 %), for testing the "64/36" surrogate fuel in Langley s Arc-Heated Scramjet Test Facility for HIFiRE engine designs. The FS results show a generally small (< 4 %) "nitric oxide enhancement" effect, relative to clean air, for up to 3 % NO (freestream Mach number up to 7 in Arc Jet testing). However, a progressively large "oxygendeficiency weakening" effect develops. For 3 % NO, a net weakening of 26 % in FS is derived for the "64/36" fuel vs. air. The corresponding net weakening for pure ethylene is 20 %. A number of practical recommendations regarding facility test effects are offered.

  5. The effect of entrance radius and film injection on wall heating in scramjet nozzles

    NASA Technical Reports Server (NTRS)

    Baker, N. R.; Northam, G. B.; Capriotti, D. P.; Stouffer, S. D.

    1992-01-01

    An experimental study has been performed with a scramjet combustor and nozzle model to determine the effect of nozzle radii on relaminarization, and the effectiveness of film injection in reducing wall heat flux. The nozzle model was installed downstream of a swept-ramp scramjet combustor model, operated at Mach 2.7 with vitiated air, simulating Mach 6-7 flight enthalpies. Results from tests conducted with three nozzle radii, varying in curvature from a sharp Prandtl-Meyer expansion to a radius of two scramjet combustor exit heights, showed little or no effect on the downstream wall heat flux, which is not in agreement with both subsonic and supersonic relaminarization correlations in the available literature. In contrast, film injection upstream of all three radii produced reductions of 70 percent in the entire nozzle model heating rate, for the nominal film-to-free stream pressure matched condition. Results from film injection, with and without pressure gradient, but without wall curvature, show that the gradients appear to stabilize the film, with the resulting reductions in wall heat transfer rates apparent far downstream of the point of injection.

  6. Testing of DLR C/C-SiC for HIFiRE 8 Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Capriotti, Diego P.; Reimer, Thomas; Kutemeyer, Marius; Smart, Michael

    2013-01-01

    Ceramic Matrix Composites (CMCs) have been proposed for hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a C/C and a C/C-SiC material system fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for HIFiRE 8, a joint Australia / AFRL hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kPa. Flat panels of the DLR C/C and the C/C-SiC were tested in the NASA Langley Direct Connect Rig (DCR) at Mach 5 and Mach 6 enthalpy for several minutes. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used in some of the tests to increase the surface temperature of the C/C-SiC panel for approximately 350degF. The final C/C-SiC panel was tested for 3 cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

  7. Re-Educating Jet-Engine-Researchers to Stay Relevant

    NASA Astrophysics Data System (ADS)

    Gal-Or, Benjamin

    2016-06-01

    To stay relevantly supported, jet-engine researchers, designers and operators should follow changing uses of small and large jet engines, especially those anticipated to be used by/in the next generation, JET-ENGINE-STEERED ("JES") fleets of jet drones but fewer, JES-Stealth-Fighter/Strike Aircraft. In addition, some diminishing returns from isolated, non-integrating, jet-engine component studies, vs. relevant, supersonic, shock waves control in fluidic-JES-side-effects on compressor stall dynamics within Integrated Propulsion Flight Control ("IPFC"), and/or mechanical JES, constitute key relevant methods that currently move to China, India, South Korea and Japan. The central roles of the jet engine as primary or backup flight controller also constitute key relevant issues, especially under post stall conditions involving induced engine-stress while participating in crash prevention or minimal path-time maneuvers to target. And when proper instructors are absent, self-study of the JES-STVS REVOLUTION is an updating must, where STVS stands for wing-engine-airframe-integrated, embedded stealthy-jet-engine-inlets, restructured engines inside Stealth, Tailless, canard-less, Thrust Vectoring IFPC Systems. Anti-terror and Airliners Super-Flight-Safety are anticipated to overcome US legislation red-tape that obstructs JES-add-on-emergency-kits-use.

  8. Rocketdyne RBCC Engine Concept Development

    NASA Technical Reports Server (NTRS)

    Ratckin, G.; Goldman, A.; Ortwerth, P.; Weisberg, S.

    1999-01-01

    Boeing Rocketdyne is pursuing the development of Rocket Based Combined Cycle (RBCC), propulsion systems as demonstrated by significant contract work in the hypersonic arena (ART, NASP, SCT, system studies) and over 12 years of steady company discretionary investment. The Rocketdyne concept is a fixed geometry integrated rocket, ramjet, scramjet which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals. seal purge gas, and closeout side attachments. Rocketdyne's experimental RBCC engine (Engine A5) was constructed under contract with the NASA Marshall Space Flight Center. Engine A5 models the complete flight engine flowpath consisting of an inlet, isolator, airbreathing combustor and nozzle. High performance rocket thrusters are integrated into the engine to enable both air-augmented rocket (AAR) and pure rocket operation. Engine A5 was tested in CASL's new FAST facility as an air-augmented rocket, a ramjet and a pure rocket. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. Rocket mode performance was above predictions. For the first time. testing also demonstrated transition from AAR operation to ramjet operation. This baseline configuration has also been shown, in previous testing, to perform well in the scramjet mode.

  9. Laser-Induced Thermal Acoustic Measurements in a Highly Back-Pressured Scramjet Isolator Model: A Research Plan

    NASA Technical Reports Server (NTRS)

    Middleton, Troy F.; Balla, Robert J.; Baurle, Robert A.; Wilson, Lloyd G.

    2008-01-01

    Under the Propulsion Discipline of NASA s Fundamental Aeronautics Program s Hypersonics Project, a test apparatus, for testing a scramjet isolator model, is being constructed at NASA's Langley Research Center. The test apparatus will incorporate a 1-inch by 2-inch by 15-inch-long scramjet isolator model supplied with 2.1 lbm/sec of unheated dry air through a Mach 2.5 converging-diverging nozzle. The planned research will incorporate progressively more challenging measurement techniques to characterize the flow field within the isolator, concluding with the application of the Laser-Induced Thermal Acoustic (LITA) measurement technique. The primary goal of this research is to use the data acquired to validate Computational Fluid Dynamics (CFD) models employed to characterize the complex flow field of a scramjet isolator. This paper describes the test apparatus being constructed, pre-test CFD simulations, and the LITA measurement technique.

  10. Testing of the Strutjet RBCC Engine

    NASA Technical Reports Server (NTRS)

    Bulman, Mel; Neill, Todd; Yam, Clement

    1999-01-01

    The testing of the Strutjet Rocket-Based Combined Cycle (RBCC) engine is described in this presentation. Since 1996 Aerojet has conducted thousands of tests on its Strutjet RBCC Engine. The Strutjet is specifically designed as an RBCC engine. The tests included: (1) Inlet, (2) Rocket Checkouts, (3) Direct Connect Mach 6 and 8, and (4) Freejet M = 0, 2.4, 4.1. The modes tested include: (1) Air Augmented Rocket (AAR), (2) Ramjet (3) Scramjet (4) Scram/Rocket (5) Ascent/Rocket. The Strutjet RBCC engine has been tested across the complete airbreathing trajectory. The presentation shows the results of the tests, through diagrams, charts and pictures.

  11. Computers and the aerospace engineer

    SciTech Connect

    Trego, L.E.

    1990-03-01

    The use of computers in aerospace for design and analysis is described, and examples of project enhancements are presented. NASA is working toward the design of a numerical test cell that will allow integrated, multidisciplinary design, analysis, and optimization of propulsion systems. It is noted that with continuing advances in computer technology, including areas such as three-dimensional computer-aided design, finite element analysis, supercomputers, and artificial intelligence, the possibilities seem limitless for the aerospace engineer. Research projects are currently underway for design and/or reconfiguration of the V-22, B-767, SCRAMJET engines, F-16, and X29A using these techniques.

  12. A forebody design technique for highly integrated bottom-mounted scramjets with application to a hypersonic research airplane

    NASA Technical Reports Server (NTRS)

    Edwards, C. L. W.

    1974-01-01

    An inviscid technique for designing forebodies which produce uniformly precompressed flows at the inlet entrance for bottom-mounted scramjets has been developed so that geometric constraints resulting from design trade-offs can be effectively evaluated. The flow fields resulting from several forebody designs generated in support of a hypersonic research airplane conceptual design study have been analyzed in detail with three-dimensional characteristics calculations to verify the uniform flow conditions. For the designs analyzed, uniform flow is maintained over a wide range of flight conditions (Mach number equals 4 to 10; angle of attack equals 6 deg to 10 deg) corresponding to scramjet operation flight envelope of the research airplane.

  13. A Computational and Experimental Investigation of a Three-Dimensional Hypersonic Scramjet Inlet Flow Field. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Holland, Scott Douglas

    1991-01-01

    A combined computational and experimental parametric study of the internal aerodynamics of a generic three dimensional sidewall compression scramjet inlet configuration was performed. The study was designed to demonstrate the utility of computational fluid dynamics as a design tool in hypersonic inlet flow fields, to provide a detailed account of the nature and structure of the internal flow interactions, and to provide a comprehensive surface property and flow field database to determine the effects of contraction ratio, cowl position, and Reynolds number on the performance of a hypersonic scramjet inlet configuration.

  14. An experimental investigation of mixing enhancement in a simulated scramjet combustor by use of swirling jets. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Kraus, Donna Karen

    1993-01-01

    It is desired to maintain supersonic flow through the combustor of supersonic airbreathing engines to reduce static temperatures and total pressure losses inherent in reducing flow to subsonic speeds. Due to the supersonic speeds through the combustor, mixing of the fuel and air must by rapid for complete combustion to occur within a reasonable streamwise distance. It was proposed that the addition of swirl to the fuel jet prior to injection might enhance the mixing of the fuel with the air. The effects of swirl on the mixing of a 30 deg wall jet into a Mach 2 flow were experimentally investigated. Swirl was introduced into the fuel stream by tangential injection into a cylindrical swirl chamber. The flow was then accelerated through a convergent-divergent nozzle with an area ratio of two, and supersonically injected into the Mach 2 flow such that the static pressure of the fuel matched the effective back pressure of the main flow. Two different cases with swirl and one without swirl were investigated, with both helium and air simulating the fuel. Rayleigh scattering was used to visualize the flow and seeding the fuel with water allowed it to be traced through the main flow. Using histograms of the pure molecular Rayleigh scattering images, the helium concentration in the jet-mixing region of the flow was monitored and found to decrease slightly with swirl, indicating better mixing. Thresholding the water-seeded images allowed the jet-mixing region to be isolated and showed a slight increase in this area with swirl. Penetration, however, was slightly less with swirl. Rescaling the data for equal mass flow rates allowed comparison for a scramjet application of a combustor with a single injector and the desire to fuel to a specified fuel-to-oxidant ratio. These results showed a substantial increase in the spreading area with swirl, an increase in the mixing occurring in this area, and slightly better penetration.

  15. Laser Diagnostics Study of Plasma Assisted Combustion for Scramjet Applications

    DTIC Science & Technology

    2011-12-01

    future. The combustion process in these engines typically involves highly turbulent reactive flow conditions, often beyond the limits of our...electric field gives rise to new electron and ion impact processes which can enhance the propagation and branching of radicals and ultimately...is generated separately and the flame is ignited as the gases pass over the plasma region. The actual oxidation process occurs further downstream

  16. Theory and Design of Flight-Vehicle Engines

    NASA Technical Reports Server (NTRS)

    Zhdanov, V. T. (Editor); Kurziner, R. I. (Editor)

    1987-01-01

    Papers are presented on such topics as the testing of aircraft engines, errors in the experimental determination of the parameters of scramjet engines, the effect of the nonuniformity of supersonic flow with shocks on friction and heat transfer in the channel of a hypersonic ramjet engine, and the selection of the basic parameters of cooled GTE turbines. Consideration is also given to the choice of optimal total wedge angle for the acceleration of aerospace vehicles, the theory of an electromagnetic-resonator engine, the dynamic characteristics of the pumps and turbines of liquid propellant rocket engines in transition regimes, and a hierarchy of mathematical models for spacecraft control engines.

  17. Conceptual Design of a Mach 8 Hypersonic Cruiser with Dorsal Engine

    NASA Astrophysics Data System (ADS)

    Murray, N.; Steelant, J.; Mack, A.

    2009-01-01

    A Mach 8 cruise concept vehicle is defined within the LAPCAT I program to explore the feasibility of antipodal travel. Different engine configurations including Rocket Ejector, Air Turbo Rocket, Ramjet and Scramjet are integrated into a waverider type vehicle and flown as part of a numerical trajectory.

  18. Numerical investigation of supersonic base flow with parallel injection. [in scramjet combustors

    NASA Technical Reports Server (NTRS)

    Sullins, G. A.; Anderson, J. D., Jr.; Drummond, J. P.

    1982-01-01

    The present investigation represents the first detailed numerical calculations of base flow with gas injection using a geometry and primary flow condition germane to a scramjet combustor. The investigation is concerned with the numerical solution of the complete two-dimensional Navier-Stokes equations for the flowfield in the vicinity of the base with parallel gas injection, taking into account the fluid dynamic aspects of this flowfield. The flow is dominated by separation, and by mixed regions of locally subsonic and supersonic flow. A comparison is conducted of flows with and without injection, and the effect of base injection on the wave patterns and shear layers of such flows is clearly established.

  19. Opposed Jet Burner Extinction Limits: Simple Mixed Hydrocarbon Scramjet Fuels vs Air

    NASA Technical Reports Server (NTRS)

    Pellett, Gerald L.; Vaden, Sarah N.; Wilson, Lloyd G.

    2007-01-01

    Opposed Jet Burner tools have been used extensively by the authors to measure Flame Strength (FS) of laminar non-premixed H2 air and simple hydrocarbon (HC) air counterflow diffusion flames at 1-atm. FS represents a strain-induced extinction limit based on air jet velocity. This paper follows AIAA-2006-5223, and provides new HC air FSs for global testing of chemical kinetics, and for characterizing idealized flameholding potentials during early scramjet-like combustion. Previous FS data included six HCs, pure and N2-diluted; and three HC-diluted H2 fuels, where FS decayed very nonlinearly as HC was added to H2, due to H-atom scavenging. This study presents FSs on mixtures of (candidate surrogate) HCs, some with very high FS ethylene. Included are four binary gaseous systems at 300 K, and a hot ternary system at approx. 600 K. The binaries are methane + ethylene, ethane + ethylene, methane + ethane, and methane + propylene. The first three also form two ternary systems. The hot ternary includes both 10.8 and 21.3 mole % vaporized n-heptane and full ranges of methane + ethylene. Normalized FS data provide accurate means of (1) validating, globally, chemical kinetics for extinction of non-premixed flames, and (2) estimating (scaling by HC) the loss of incipient flameholding in scramjet combustors. The n-heptane is part of a proposed baseline simulant (10 mole % with 30% methane + 60% ethylene) that mimics the ignition of endothermically cracked JP-7 like kerosene fuel, as suggested by Colket and Spadaccini in 2001 in their shock tube Scramjet Fuels Autoignition Study. Presently, we use FS to gauge idealized flameholding, and define HC surrogates. First, FS was characterized for hot nheptane + methane + ethylene; then a hot 36 mole % methane + 64% ethylene surrogate was defined that mimics FS of the baseline simulant system. A similar hot ethane + ethylene surrogate can also be defined, but it has lower vapor pressure at 300 K, and thus exhibits reduced gaseous

  20. Laser-Induced Thermal Acoustics Theory and Expected Experimental Errors when Applied to a Scramjet Isolator Model

    NASA Technical Reports Server (NTRS)

    Middleton, Troy F.; Balla, Robert Jeffrey; Baurle, Robert A.; Wilson, Lloyd G.

    2011-01-01

    A scramjet isolator model test apparatus is being assembled in the Isolator Dynamics Research Lab (IDRL) at the NASA Langley Research Center in Hampton, Virginia. The test apparatus is designed to support multiple measurement techniques for investigating the flow field in a scramjet isolator model. The test section is 1-inch high by 2-inch wide by 24-inch long and simulates a scramjet isolator with an aspect ratio of two. Unheated, dry air at a constant stagnation pressure and temperature is delivered to the isolator test section through a Mach 2.5 planar nozzle. The isolator test section is mechanically back-pressured to contain the resulting shock train within the 24-inch isolator length and supports temperature, static pressure, and high frequency pressure measurements at the wall. Additionally, nonintrusive methods including laser-induced thermal acoustics (LITA), spontaneous Raman scattering, particle image velocimetry, and schlieren imaging are being incorporated to measure off-wall fluid dynamic, thermodynamic, and transport properties of the flow field. Interchangeable glass and metallic sidewalls and optical access appendages permit making multiple measurements simultaneously. The measurements will be used to calibrate computational fluid dynamics turbulence models and characterize the back-pressured flow of a scramjet isolator. This paper describes the test apparatus, including the optical access appendages; the physics of the LITA method; and estimates of LITA measurement uncertainty for measurements of the speed of sound and temperature.

  1. Two-dimensional viscous flow computations of hypersonic scramjet nozzle flowfields at design and off-design conditions

    NASA Technical Reports Server (NTRS)

    Harloff, G. J.; Lai, H. T.; Nelson, E. S.

    1988-01-01

    The PARC2D code has been selected to analyze the flowfields of a representative hypersonic scramjet nozzle over a range of flight conditions from Mach 3 to 20. The flowfields, wall pressures, wall skin friction values, heat transfer values and overall nozzle performance are presented.

  2. Strutjet RBCC Engine Performance

    NASA Technical Reports Server (NTRS)

    Bulman, Mel; Neill, Todd; Yam, Clement

    1999-01-01

    In the past two years Strutjet Rocket Based Combined Cycle (RBCC) engine has been tested extensively under the Advanced Reusable Technology (ART) contract from NASA MSFC. RBCC Engines combine the high thrust to weight of the rocket with the high efficiency of the ramjet engine. This propulsion system has the potential to reduce the cost of launching payloads to orbit by up to a factor of 100. In the ART program we have conducted over 100 hot fire tests. The propellants have been hydrogen and oxygen. The Modes tested have included the Air Augmented Rocket (AAR) from M = 0 to 2.4, the Ramjet at M = 2.4 & 6, Scramjet at M = 6 & 8, Scram/Rocket at Mach 8 and Ascent Rocket in Vacuum. This invited paper will present an overview of these test results and plans for future development of this propulsion cycle.

  3. Dual-Pump CARS Measurements in the University of Virginia's Dual-Mode Scramjet: Configuration "C"

    NASA Technical Reports Server (NTRS)

    Cutler, Andrew D.; Magnotti, Gaetano; Cantu, Luca; Gallo, Emanuela; Danehy, Paul M.; Rockwell, Robert; Goyne, Christopher; McDaniel, James

    2013-01-01

    Measurements have been conducted at the University of Virginia Supersonic Combustion Facility in configuration C of the dual-mode scramjet. This is a continuation of previously published works on configuration A. The scramjet is hydrogen fueled and operated at two equivalence ratios, one representative of the scram mode and the other of the ram mode. Dual-pump CARS was used to acquire the mole fractions of the major species as well as the rotational and vibrational temperatures of N2. Developments in methods and uncertainties in fitting CARS spectra for vibrational temperature are discussed. Mean quantities and the standard deviation of the turbulent fluctuations at multiple planes in the flow path are presented. In the scram case the combustion of fuel is completed before the end of the measurement domain, while for the ram case the measurement domain extends into the region where the flow is accelerating and combustion is almost completed. Higher vibrational than rotational temperature is observed in those parts of the hot combustion plume where there is substantial H2 (and hence chemical reaction) present.

  4. Evaluation of scramjet nozzle configurations and film cooling for reduction of wall heating

    NASA Technical Reports Server (NTRS)

    Baker, N. R.; Northam, G. B.; Stouffer, S. D.; Capriotti, D. P.

    1993-01-01

    Experiments on relaminarization, or reverse transition of turbulent flow, have been conducted with a scramjet combustor and nozzle model. The Mach-2.7 direct-connect scramjet combustor employed swept-ramp fuel injectors, and the entering flow simulated Mach-6 to Mach-7 flight conditions. Tests were also conducted without combustor fuel to produce a higher nozzle entrance Mach number and lower pressures. Additional tests were conducted to evaluate the effectiveness of film cooling on the cowl region. At each test condition studied, the entrance radius had no significant effect on the measured downstream wall heating rates. Analysis of the data when fuel was injected into the combustor shows good agreement with turbulent theory; for the tests without fuel injection and combustion, the data downstream of the nozzle entrance radius are intermediate to the calculated turbulent and laminar solutions. Results of film-cooling tests, conducted with both hydrogen and nitrogen injectants, showed that, per unit mass, the hydrogen was approximately three to four times more effective than nitrogen in reducing the downstream wall heating rates.

  5. Thermal behavior in the cracking reaction zone of scramjet cooling channels at different channel aspect ratios

    NASA Astrophysics Data System (ADS)

    Zhang, Silong; Feng, Yu; Jiang, Yuguang; Qin, Jiang; Bao, Wen; Han, Jiecai; Haidn, Oskar J.

    2016-10-01

    To study the thermal behavior in the cracking reaction zone of regeneratively cooled scramjet cooling channels at different aspect ratios, 3-D model of fuel flow in terms of the fuel's real properties and cracking reaction is built and validated through experiments. The whole cooling channel is divided into non-cracking and cracking reaction zones. Only the cracking reaction zone is studied in this article. The simulation results indicate that the fuel conversion presents a similar distribution with temperature because the fuel conversion in scramjet cooling channels is co-decided by the temperature and velocity but the temperature plays the dominate role. For the cases given in this paper, increasing the channel aspect ratio will increase the pressure drop and it is not beneficial for reducing the wall temperature because of the much severer thermal stratification, larger conversion non-uniformity, the corresponding M-shape velocity profile which will cause local heat transfer deterioration and the decreased chemical heat absorption. And the decreased chemical heat absorption caused by stronger temperature and conversion non-uniformities is bad for the utilization of chemical heat sink, chemical recuperation process and the ignition performance.

  6. Effects of compression and expansion ramp fuel injector configuration on scramjet combustion and heat transfer

    NASA Technical Reports Server (NTRS)

    Stouffer, Scott D.; Baker, N. R.; Capriotti, D. P.; Northam, G. B.

    1993-01-01

    A scramjet combustor with four wall-ramp injectors containing Mach-1.7 fuel jets in the base of the ramps was investigated experimentally. During the test program, two swept ramp injector designs were evaluated. One swept-ramp model had 10-deg compression-ramps and the other had 10-deg expansion cavities between flush wall ramps. The scramjet combustor model was instrumented with pressure taps and heat-flux gages. The pressure measurements indicated that both injector configurations were effective in promoting mixing and combustion. Autoignition occurred for the compression-ramp injectors, and the fuel began to burn immediately downstream of the injectors. In tests of the expansion ramps, a pilot was required to ignite the fuel, and the fuel did not burn for a distance of at least two gaps downstream of the injectors. Once initiated, combustion was rapid in this configuration. Heat transfer measurements showed that the heat flux differed greatly both across the width of the combustor and along the length of the combustor.

  7. Analytical study of the hydrogen-air reaction mechanism with application to scramjet combustion

    SciTech Connect

    Jachimowski, C.J.

    1988-02-01

    A chemical kinetic mechanism for the combustion of hydrogen has been assembled and optimized by comparing the observed behavior as determined in shock tube and flame studies with that predicted by the mechanism. The reactions contained in the mechanism reflect the current state of knowledge of the chemistry of the hydrogen/air system, and the assigned rate coefficients are consistent with accepted values. It was determined that the mechanism is capable of satisfactorily reproducing the experimental results for a range of conditions relevant to scramjet combustion. Calculations made with the reaction mechanism for representative scramjet combustor conditions at Mach 8, 16, and 25 showed that chemical kinetic effects can be important and that combustor models which use nonequilibrium chemistry should be used in preference to models that assume equilibrium chemistry. For the conditions examined the results also showed the importance of including the HO/sub 2/ chemistry in the mechanism. For Mach numbers less than 16, the studies suggest that an ignition source will most likely be required to overcome slow ignition chemistry. At Mach 25, the initial temperature and pressure was high enough that ignition was rapid and the presence of an ignition source did not significantly affect reaction rates.

  8. Numerically simulated comparative performance of a scramjet and shcramjet at Mach 11

    NASA Astrophysics Data System (ADS)

    Chan, Jonathan

    This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and altitude of 34.5 km. The design includes a Prandtl-Meyer compression inlet, cantilevered ramp fuel injectors, a shock-inducing wedge and a divergent nozzle. Numerical studies are undertaken using the WARP code that solves the three-dimensional Favre-averaged Navier-Stokes equations closed by the Wilcox k-o turbulence model and the Jachimowski H2/air chemical kinetics model. Studies of fuel injection properties, mixing duct length, combustor wedge and nozzle geometry are completed to maximize the overall performance of the vehicle. The final shcramjet configuration generates a specific impulse of 1110 s. A comparison is undertaken with a scramjet vehicle at identical flight conditions and using many of the same components. The comparable scramjet generates a higher specific impulse of 1450 s although it is significantly larger and therefore heavier.

  9. Internal aerodynamics of a generic three-dimensional scramjet inlet at Mach 10

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1995-01-01

    A combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall compression scramjet inlet configuration at Mach 10 has been performed. The study was designed to demonstrate the utility of computational fluid dynamics as a design tool in hypersonic inlet flow fields, to provide a detailed account of the nature and structure of the internal flow interactions, and to provide a comprehensive surface property and flow field database to determine the effects of contraction ratio, cowl position, and Reynolds number on the performance of a hypersonic scramjet inlet configuration. The work proceeded in several phases: the initial inviscid assessment of the internal shock structure, the preliminary computational parametric study, the coupling of the optimized configuration with the physical limitations of the facility, the wind tunnel blockage assessment, and the computational and experimental parametric study of the final configuration. Good agreement between computation and experimentation was observed in the magnitude and location of the interactions, particularly for weakly interacting flow fields. Large-scale forward separations resulted when the interaction strength was increased by increasing the contraction ratio or decreasing the Reynolds number.

  10. Functional mathematical model of a hydrogen-driven combustion chamber for a scramjet

    NASA Astrophysics Data System (ADS)

    Latypov, A. F.

    2015-09-01

    A functional mathematical model of a hydrogen-driven combustion chamber for a scramjet is described. The model is constructed with the use of one-dimensional steady gas-dynamic equations and parametrization of the channel configuration and the governing parameters (fuel injection into the flow, fuel burnout along the channel, dissipation of kinetic energy, removal of some part of energy generated by gases for modeling cooling of channel walls by the fuel) with allowance for real thermophysical properties of the gases. Through parametric calculations, it is found that fuel injection in three cross sections of the channel consisting of segments with weak and strong expansion ensures a supersonic velocity of combustion products in the range of free-stream Mach numbers M∞ = 6-12. It is demonstrated that the angle between the velocity vectors of the gaseous hydrogen flow and the main gas flow can be fairly large in the case of distributed injection of the fuel. This allows effective control of the mixing process. It is proposed to use the exergy of combustion products as a criterion of the efficiency of heat supply in the combustion chamber. Based on the calculated values of exergy, the critical free-stream Mach number that still allows scramjet operation is estimated.

  11. An analytical study of the hydrogen-air reaction mechanism with application to scramjet combustion

    NASA Technical Reports Server (NTRS)

    Jachimowski, Casimir J.

    1988-01-01

    A chemical kinetic mechanism for the combustion of hydrogen has been assembled and optimized by comparing the observed behavior as determined in shock tube and flame studies with that predicted by the mechanism. The reactions contained in the mechanism reflect the current state of knowledge of the chemistry of the hydrogen/air system, and the assigned rate coefficients are consistent with accepted values. It was determined that the mechanism is capable of satisfactorily reproducing the experimental results for a range of conditions relevant to scramjet combustion. Calculations made with the reaction mechanism for representative scramjet combustor conditions at Mach 8, 16, and 25 showed that chemical kinetic effects can be important and that combustor models which use nonequilibrium chemistry should be used in preference to models that assume equilibrium chemistry. For the conditions examined the results also showed the importance of including the HO2 chemistry in the mechanism. For Mach numbers less than 16, the studies suggest that an ignition source will most likely be required to overcome slow ignition chemistry. At Mach 25, the initial temperature and pressure was high enough that ignition was rapid and the presence of an ignition source did not significantly affect reaction rates.

  12. Hyper-X Wind Tunnel Program

    NASA Technical Reports Server (NTRS)

    McClinton, C. R.; Holland, S. D.; Rock, K. E.; Engelund, W. C.; Voland, R. T.; Huebner, L. D.; Roger, R. C.

    1998-01-01

    This paper provides an overview of NASA's focused hypersonic technology program, called the Hyper-X Program. The Hyper-X Program, a joint NASA Langley and Dryden program, is designed to move hypersonic, air breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated scramjet (supersonic combustion ramjet) propulsion system at true flight conditions and the first opportunity for flight validation of experimental wind tunnel, numerical and analytical methods used for design of these vehicles. A substantial portion of the program is experimentally based, both for database development and performance validation. The program is now concentrating on Mach 7 vehicle development, verification and validation and flight test risk reduction. This paper concentrates on the aerodynamic and propulsion experimental programs. Wind tunnel testing of the flight engine and complete airframe integrated scramjet configuration flow-path is expected in 1998 and 1999, respectively, and flight test is planned for 2000.

  13. Current MBDA R&T Effort on Ram/Scramjet and Detonation Wave Engine

    DTIC Science & Technology

    2006-05-01

    by a ceramic composite heat shield protecting from erosion and by a layer of insulation material were tested and solutions, able to sustain low...knowledge and preserve human and material investments in spite of the lack of new National or European R&T program [12] : • JAPHAR program (ONERA and DLR...stored in a high pressure bottle generally placed inside the tank. The transmission of the pressure to the fuel is done by an elastomeric bladder. A

  14. The enhancement of the mixing and combustion processes in supersonic flow applied to scramjet engine

    SciTech Connect

    Kopchenov, V.I.; Lomkov, K.E. )

    1992-07-01

    The Reynolds averaged parabolized Navier-Stokes equations are employed for the numerical study of turbulent mixing and combustion of a supersonic hydrogen jet in a supersonic airflow. A one-equation differential turbulence model is utilized. The simplified flame sheet model is employed for the numerical simulation of the supersonic combustion. 24 refs.

  15. Back-pressure Effect on Shock-Train Location in a Scramjet Engine Isolator

    DTIC Science & Technology

    2010-03-01

    window the pressure causes. The book, Roak’s formulas for stress and strain written by Warren C. Yong and Richard G. Budynas , was used as a reference...Warren C. and Budynas , Richard G. Roak’s formula for stress and strain. 7th Edition. s.l. : McGraw-Hill, 2002. 176 Vita Captain Richard I. Paek

  16. Visualization of hydrogen injection in a scramjet engine by simultaneous PLIF imaging and laser holographic imaging

    NASA Technical Reports Server (NTRS)

    Anderson, Robert C.; Trucco, Richard E.; Rubin, L. F.; Swain, D. M.

    1992-01-01

    Flowfield characterization has been accomplished for several fuel injector configurations using simultaneous planar laser induced fluorescence (PLIF) and laser holographic imaging (LHI). The experiments were carried out in the GASL-NASA HYPULSE real gas expansion tube facility, a pulsed facility with steady test times of about 350 microsec. The tests were done at simulated Mach numbers 13.5 and 17. The focus of this paper is on the measurement technologies used and their application in a research facility. The HYPULSE facility, the models used for the experiments, and the setup for the LHI and PLIF measurements are described. Measurement challenges and solutions are discussed. Results are presented for experiments with several fuel injector configurations and several equivalence ratios.

  17. Application of Modern Design of Experiments to CARS Thermometry in a Model Scramjet Engine

    NASA Technical Reports Server (NTRS)

    Danehy, P. M.; DeLoach, R.; Cutler, A. D.

    2002-01-01

    We have applied formal experiment design and analysis to optimize the measurement of temperature in a supersonic combustor at NASA Langley Research Center. We used the coherent anti-Stokes Raman spectroscopy (CARS) technique to map the temperature distribution in the flowfield downstream of an 1160 K, Mach 2 freestream into which supersonic hydrogen fuel is injected at an angle of 30 degrees. CARS thermometry is inherently a single-point measurement technique; it was used to map thc flow by translating the measurement volume through the flowfield. The method known as "Modern Design of Experiments" (MDOE) was used to estimate the data volume required, design the test matrix, perform the experiment and analyze the resulting data. MDOE allowed us to match the volume of data acquired to the precision requirements of the customer. Furthermore, one aspect of MDOE, known as response surface methodology, allowed us to develop precise maps of the flowfield temperature, allowing interpolation between measurement points. An analytic function in two spatial variables was fit to the data from a single measurement plane. Fitting with a Cosine Series Bivariate Function allowed the mean temperature to be mapped with 95% confidence interval half-widths of +/- 30 Kelvin, comfortably meeting the confidence of +/- 50 Kelvin specified prior to performing the experiments. We estimate that applying MDOE to the present experiment saved a factor of 5 in data volume acquired, compared to experiments executed in the traditional manner. Furthermore, the precision requirements could have been met with less than half the data acquired.

  18. Review of Air Vitiation Effects on Scramjet Ignition and Flameholding Combustion Processes

    NASA Technical Reports Server (NTRS)

    Pellett, G. L.; Bruno, Claudio; Chinitz, W.

    2002-01-01

    This paper offers a detailed review and analysis of more than 100 papers on the physics and chemistry of scramjet ignition and flameholding combustion processes, and the known effects of air vitiation on these processes. The paper attempts to explain vitiation effects in terms of known chemical kinetics and flame propagation phenomena. Scaling methodology is also examined, and a highly simplified Damkoehler scaling technique based on OH radical production/destruction is developed to extrapolate ground test results, affected by vitiation, to flight testing conditions. The long term goal of this effort is to help provide effective means for extrapolating ground test data to flight, and thus to reduce the time and expense of both ground and flight testing.

  19. The effect of kerosene injection on ignition probability of local ignition in a scramjet combustor

    NASA Astrophysics Data System (ADS)

    Bao, Heng; Zhou, Jin; Pan, Yu

    2017-03-01

    The spark ignition of kerosene is investigated in a scramjet combustor with a flight condition of Ma 4, 17 km. Based plentiful of experimental data, the ignition probabilities of the local ignition have been acquired for different injection setups. The ignition probability distributions show that the injection pressure and injection location have a distinct effect on spark ignition. The injection pressure has both upper and lower limit for local ignition. Generally, the larger mass flow rate will reduce the ignition probability. The ignition position also affects the ignition near the lower pressure limit. The reason is supposed to be the cavity swallow effect on upstream jet spray near the leading edge, which will make the cavity fuel rich. The corner recirculation zone near the front wall of the cavity plays a significant role in the stabilization of local flame.

  20. A numerical procedure for the parametric optimization of three dimensional scramjet nozzles

    NASA Technical Reports Server (NTRS)

    Dash, S.; Delguidice, P.; Kalben, P.

    1973-01-01

    A numerical procedure permitting the rapid determination of the internal performance of a class of scramjet nozzle configurations is presented. The approach developed is based on the construction of quasi two dimensional simple wave networks, where lateral expansion effects are incorporated by one dimensional approximations. A numerical procedure following this approach has has been developed and results obtained are highly comparable to those obtained employing a characteristic procedure. The numerical program developed permits the parametric variation of cowl length, turning angles on the cowl and vehicle undersurface and lateral expansion and is subject to fixed constraints such as the vehicle length and nozzle exit height. The program requires uniform initial conditions at the burner exit station and yields the location of all predominant wave zones, accounting for lateral expansion effects. In addition, the program yields the detailed pressure distribution on the cowl and vehicle undersurface and calculates the nozzle thrust, lift and pitching moment.

  1. CFD prediction of the reacting flow field inside a subscale scramjet combustor

    NASA Technical Reports Server (NTRS)

    Chitsomboon, T.; Northam, G. B.; Rogers, R. C.; Diskin, G. S.

    1988-01-01

    A three-dimensional, Reynolds-averaged Navier-Stokes CFD code has been used to calculate the reacting flowfield inside a hydrogen-fueled, subscale scramjet combustor. Pilot fuel was injected transversely upstream of the combustor and the primary fuel was injected transversely downstream of a backward facing step. A finite rate combustion model with two-step kinetics was used. The CFD code used the explicit MacCormack algorithm with point-implicit treatment of the chemistry source terms. Turbulent mixing of the jets with the airstream was simulated by a simple mixing length scheme, whereas near wall turbulence was accounted for by the Baldwin-Lomax model. Computed results were compared with experimental wall pressure measurements.

  2. Computational models for the analysis/design of hypersonic scramjet components. I - Combustor and nozzle models

    NASA Technical Reports Server (NTRS)

    Dash, S. M.; Sinha, N.; Wolf, D. E.; York, B. J.

    1986-01-01

    An overview of computational models developed for the complete, design-oriented analysis of a scramjet propulsion system is provided. The modular approach taken involves the use of different PNS models to analyze the individual propulsion system components. The external compression and internal inlet flowfields are analyzed by the SCRAMP and SCRINT components discussed in Part II of this paper. The combustor is analyzed by the SCORCH code which is based upon SPLITP PNS pressure-split methodology formulated by Dash and Sinha. The nozzle is analyzed by the SCHNOZ code which is based upon SCIPVIS PNS shock-capturing methodology formulated by Dash and Wolf. The current status of these models, previous developments leading to this status, and, progress towards future hybrid and 3D versions are discussed in this paper.

  3. Experimental Investigation of Reacting Flow Characteristics in a Dual-Mode Scramjet Combustor

    NASA Astrophysics Data System (ADS)

    Shi, Deyong; Song, Wenyan; Ye, Jingfeng; Tao, Bo; Wang, Yanhua

    2016-06-01

    In this work, a hydrogen-fueled dual-mode scramjet combustor was investigated experimentally. Clean and dry air was supplied to the combustor through a Mach 2 nozzle with a total temperature of 800 K and a total pressure of 800 kPa. The high enthalpy air was provided by an electricity resistance heater. Room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted. Except wall pressure profiles, velocity and temperature profiles in and at exit of the combustor were also measured using hydroxyl tagging velocimetry (HTV) and tunable diode laser absorption spectroscopy (TDLAS), respectively. Results showed that combustion occurred mainly at the bottom side of the combustor. And there were also an extreme disparity of the velocity and temperature profiles along the Y-direction, i.e. the transverse direction.

  4. Large Eddy Simulation Of The Fuel Injection In Scramjet Combustion Chambers

    NASA Astrophysics Data System (ADS)

    Rana, Z. A.; Thornber, R. J. R.; Drikakis, D.

    2011-05-01

    Fuel injection in a typical scramjet combustion chamber is a very challenging flow to characterise either in experimental or computational studies. It involves multi- species compressible turbulent features with complex coherent flow structures arising as a result of sonic fuel injection transverse or inclined to the free-stream super- sonic flow. In this paper the fuel injection in the HyShot- II combustion chamber is studied using an Implicit LES method employing a modified very high order accurate numerical method. To gain accurate mean inflow boundary conditions, a thermally perfect gas formulation has been employed in preliminary simulations of the inlet ramp and cowl configuration. The results of these simulations are presented and validated against wind tunnel data.

  5. Multiple-scale turbulence modeling of boundary layer flows for scramjet applications

    NASA Technical Reports Server (NTRS)

    Fabris, G.; Harsha, P. T.; Edelman, R. B.

    1981-01-01

    As part of an investigation into the application of turbulence models to the computation of flows in advanced scramjet combustors, the multiple-scale turbulence model was applied to a variety of flowfield predictions. The model appears to have a potential for improved predictions in a variety of areas relevant to combustor problems. This potential exists because of the partition of the turbulence energy spectrum that is the major feature of the model and which allows the turbulence energy dissipation rate to be out of phase with turbulent energy production. The computations were made using a consistent method of generating experimentally unavailable initial conditions. An appreciable overall improvement in the generality of the predictions is observed, as compared to those of the basic two-equation turbulence model. A Mach number-related correction is found to be necessary to satisfactorily predict the spreading rate of the supersonic jet and mixing layer.

  6. An Experimental Parametric Study of Geometric, Reynolds Number, and Ratio of Specific Heats Effects in Three-Dimensional Sidewall Compression Scramjet Inlets at Mach 6

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.; Murphy, Kelly J.

    1993-01-01

    Since mission profiles for airbreathing hypersonic vehicles such as the National Aero-Space Plane include single-stage-to-orbit requirements, real gas effects may become important with respect to engine performance. The effects of the decrease in the ratio of specific heats have been investigated in generic three-dimensional sidewall compression scramjet inlets with leading-edge sweep angles of 30 and 70 degrees. The effects of a decrease in ratio of specific heats were seen by comparing data from two facilities in two test gases: in the Langley Mach 6 CF4 Tunnel in tetrafluoromethane (where gamma=1.22) and in the Langley 15-Inch Mach 6 Air Tunnel in perfect gas air (where gamma=1.4). In addition to the simulated real gas effects, the parametric effects of cowl position, contraction ratio, leading-edge sweep, and Reynolds number were investigated in the 15-Inch Mach 6 Air Tunnel. The models were instrumented with a total of 45 static pressure orifices distributed on the sidewalls and baseplate. Surface streamline patterns were examined via oil flow, and schlieren videos were made of the external flow field. The results of these tests have significant implications to ground based testing of inlets in facilities which do not operate at flight enthalpies.

  7. Scramjet fuel injector design parameters and considerations: Development of a two-dimensional tangential fuel injector with constant pressure at the flame

    NASA Technical Reports Server (NTRS)

    Agnone, A. M.

    1972-01-01

    The factors affecting a tangential fuel injector design for scramjet operation are reviewed and their effect on the efficiency of the supersonic combustion process is evaluated using both experimental data and theoretical predictions. A description of the physical problem of supersonic combustion and method of analysis is followed by a presentation and evaluation of some standard and exotic types of fuel injectors. Engineering fuel injector design criteria and hydrogen ignition schemes are presented along with a cursory review of available experimental data. A two-dimensional tangential fuel injector design is developed using analyses as a guide in evaluating the effects on the combustion process of various initial and boundary conditions including splitter plate thickness, injector wall temperature, pressure gradients, etc. The fuel injector wall geometry is shaped so as to maintain approximately constant pressure at the flame as required by a cycle analysis. A viscous characteristics program which accounts for lateral as well as axial pressure variations due to the mixing and combustion process is used in determining the wall geometry.

  8. Design Of A SCRAMJET Nozzle With Streamline Tracing Technique And Reference Temerature Methode

    NASA Astrophysics Data System (ADS)

    Riehmer, J.; Gulhan, A.

    2011-05-01

    This study presents a method to find an optimal shape of a three-dimensional supersonic nozzle for a rectangular scramjet combustion chamber with rounded edges by taking into account the skin friction effects. The geometric and flow constraints are defined within the German DFG GRK 1095/2 project and the designed nozzle will be part of a scramjet demonstrator configuration [1]. The nozzle inlet conditions are mean values of the combustion chamber exit conditions with the assumption of a constant specific heat ratio. To generate the shape of the nozzle a streamline tracing technique is applied to an axis-symmetric flow field calculated by the Method of Characteristics (MOC). Skin friction in relatively high pressure supersonic flow from the combustion chamber is very dominant and cannot be neglected in the design process. Therefore the skin friction is calculated using the Reference Temperature Method (RTM) and used for the determination of the thrust and moment vectors. This allows considering viscous effects without boundary layer calculations. With this approach an optimal truncated ideal nozzle contour which yields the geometric constraints can be derived. For the validation of this method comparative calculations have been carried out with the DLR code TAU on an exemplary axis-symmetric supersonic nozzle for different flow conditions. Results showed a good agreement. Finally for the three-dimensional nozzle the analytical solution for the inviscous and viscous case provided comparable data like TAU simulations. Further simplifications of the approach for an efficient three-dimensional nozzle design will be addressed in the paper.

  9. Numerical Investigation of Dual-Mode Scramjet Combustor with Large Upstream Interaction

    NASA Technical Reports Server (NTRS)

    Mohieldin, T. O.; Tiwari, S. N.; Reubush, David E. (Technical Monitor)

    2004-01-01

    Dual-mode scramjet combustor configuration with significant upstream interaction is investigated numerically, The possibility of scaling the domain to accelerate the convergence and reduce the computational time is explored. The supersonic combustor configuration was selected to provide an understanding of key features of upstream interaction and to identify physical and numerical issues relating to modeling of dual-mode configurations. The numerical analysis was performed with vitiated air at freestream Math number of 2.5 using hydrogen as the sonic injectant. Results are presented for two-dimensional models and a three-dimensional jet-to-jet symmetric geometry. Comparisons are made with experimental results. Two-dimensional and three-dimensional results show substantial oblique shock train reaching upstream of the fuel injectors. Flow characteristics slow numerical convergence, while the upstream interaction slowly increases with further iterations. As the flow field develops, the symmetric assumption breaks down. A large separation zone develops and extends further upstream of the step. This asymmetric flow structure is not seen in the experimental data. Results obtained using a sub-scale domain (both two-dimensional and three-dimensional) qualitatively recover the flow physics obtained from full-scale simulations. All results show that numerical modeling using a scaled geometry provides good agreement with full-scale numerical results and experimental results for this configuration. This study supports the argument that numerical scaling is useful in simulating dual-mode scramjet combustor flowfields and could provide an excellent convergence acceleration technique for dual-mode simulations.

  10. Flow Visualization of a Scramjet Inlet - Isolator Model in Supersonic Flow

    NASA Astrophysics Data System (ADS)

    Seckin, S.; Yuceil, K. B.

    2013-04-01

    Understanding the physical mechanisms and having insight to the complex flowfield involving unstart phenomena in supersonic inlets has gained considerable attention especially in the area of scramjet inlet/isolator aerothermodynamics. In this study, Schlieren visualization and computational analysis of shock wave structures in ramjet/scramjet inlet/isolator models in supersonic flow have been performed. Experiments were performed in the supersonic wind tunnel at the Trisonic Research Laboratory in Istanbul Technical University. The test section floor and the existing mechanism underneath have been modified to be able to mount the designed inlet/isolator model on the floor of the test section. The inlet/isolator model with a 12- degree compression ramp is investigated at Mach 2 both computationally and experimentally. Computations were performed using Star-CCM+ software to investigate shock wave structures in and around the three dimensional inlet/isolator model as mounted on the test section floor as a guide for designing the experimental model. In the results, the effects of shock wave - boundary layer interactions with flow separations with were observed. Ensemble average of the density distributions on a series of planes from one side wall to the other from the CFD results agreed well with the Schlieren images obtained experimentally. The structure of the shock waves and angles obtained from the Schlieren images agree quite well with those obtained from the CFD results. The effects of lambda-shock formations which indicate possible boundary layer separations, reflections of shock waves, and shock wave - boundary layer interactions on inlet unstart phenomena have been discussed. In order to investigate inlet unstart mechanism further, different experimental setups have been suggested for future work.

  11. A forebody design technique for highly integrated bottom-mounted scramjets with application to a hypersonic research airplane

    NASA Technical Reports Server (NTRS)

    Edwards, C. L. W.

    1976-01-01

    A rapid and simple inviscid technique for designing forebodies which produce uniformly precompressed flows at the entrance for bottom-mounted scramjets was developed so that geometric constraints resulting from design trade-offs can be effectively evaluated. The flow fields resulting from several forebody designs generated in support of a conceptual design for a hypersonic research airplane have been analyzed in detail. Three-dimensional characteristics calculations were used to verify the uniform flow conditions. For the designs analyzed, uniform flow was maintained over a wide range of flight conditions (Mach numbers from 4 to 10; angles of attack of 6 and 10 deg) corresponding to the scramjet operation flight envelope of the research airplane.

  12. Laser-Based Measurements of OH, Temperature, and Water Vapor Concentration in a Hydrocarbon-Fueled Scramjet (POSTPRINT)

    DTIC Science & Technology

    2008-07-01

    its cowl wall. The combustor had over 200 pressure taps instrumented on all four walls, including measurements at the combustor exit and in the base...connect hydrocarbon-fueled scramjet combustor . Planar laser- induced fluorescence (PLIF) of the OH radical is used to examine the flame structure...within the combustor . Tunable diode laser- based absorption spectroscopy (TDLAS) is used to measure water vapor concentration and static temperature near

  13. Testing of DLR C/C-SiC and C/C for HIFiRE 8 Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Capriotti, Diego P.; Reimer, Thomas; Kutemeyer, Marius; Smart, Michael K.

    2014-01-01

    Ceramic Matrix Composites (CMCs) have been proposed for use as lightweight hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a carbon/carbon (C/C) and a carbon/carbon-silicon carbide (C/C-SiC) material fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for Hypersonic International Flight Research Experimentation (HIFiRE) 8, a joint Australia / Air Force Research Laboratory hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kilopascals. Flat panels of the DLR C/C and C/C-SiC materials were installed downstream of a hydrogen-fueled, dual-mode scramjet combustor and tested for several minutes at conditions simulating flight at Mach 5 and Mach 6. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used during some of the tests to reduce heat loss from the back surface and thus increase the heated surface temperature of the C/C-SiC panel approximately 177 C (350 F). The final C/C-SiC panel was tested for three cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

  14. Scramjet Nozzles

    DTIC Science & Technology

    2010-09-01

    integration et gestion thermique) 14. ABSTRACT The lecture is given in four parts, each being a step in the process of nozzle design, and within each part...exhaust is directed downward at θ3, turning the lee airflow through the same angle and resulting in a pressure p3 that is lower than ambient and...below ambient proved entirely incompatible with any 3D implementation of the ideal 2D flow field, of which we could conceive. We were left with

  15. Scramjet injector

    SciTech Connect

    Bulman, M.J.

    1993-06-22

    A fuel injector mounted flush to the wall of a combustor through which air flows at a predetermined air pressure in a direction defined as the air flow direction, the fuel injector is described comprising: a generally elongated body including a surface substantially flush with the wall, the surface having a proximate end and a distal end aligned with the air flow direction such that the air flow direction extends from the proximate end to the distal end; at least one fuel inlet port connected to the generally elongated body; a single elongated exit slot having a longitudinal axis parallel to the air flow direction, the elongated exit slot having an exit slot cross section perpendicular to the longitudinal axis, the exit slot cross section having a width parallel to the surface and a depth perpendicular to the surface, the width and depth increasing in the air flow direction; and at least one throat in the generally elongated body communicating the fuel inlet port with the single elongated exit slot, wherein the single elongated exit slot and the throat are configured to create and maintain a substantially streamlined supersonic fuel jet plume shape having a local exit pressure substantially equal to the predetermined air pressure.

  16. Numerical Simulation of High-Speed Combustion Processes in Scramjet Configurations

    NASA Astrophysics Data System (ADS)

    Potturi, Amarnatha Sarma

    Flows through scramjet configurations are simulated using hybrid large-eddy simulation / Reynolds-averaged Navier-Stokes techniques. Present study is performed in three parts: parametric studies to determine the sensitivities of the predictions to modeling and algorithmic variations; formulation, implementation, and testing of several subgrid closures aimed at modeling filtered species production rates, which account for turbulence-chemistry interactions in a finite rate chemistry large-eddy simulation framework; and as a final assessment of the complete methodology, cavity-stabilized ethylene combustion is simulated. Throughout the present study, emphasis is placed on characterizing facility-specific effects, since they can have a significant influence on the numerical solution. In Part One, non-reactive and reactive flows through a model scramjet combustor with a wedge shaped injector are simulated. Different grids, flux reconstruction methods, reaction mechanisms, and inflow boundary conditions are used. To enhance fuel-air mixing, a synthetic eddy method is used to generate turbulence in the injector boundary layers and the hydrogen jets. The results show that in all the cases a lifted flame is predicted with varying standoff distances, heat releases, and shapes. In Part Two, the subgrid closures for modeling the filtered species production rates are tested on two different scramjet configurations with fundamentally different flow patterns and flame structures, one with the wedge shaped injector placed at the center of the combustor section (first, used in Part One), another with a three-dimensional ramp injector located on the upper wall of the combustor section (second). While the impact of these closures on the flow through the first configuration is insignificant, they have a more pronounced effect on the flow through the second configuration. Error analysis and performance quantification of these closures reveal that, relative to a baseline model, two of

  17. Hypersonic engine seal development at NASA Lewis Research Center

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.

    1994-01-01

    NASA Lewis Research Center is developing advanced seal concepts and sealing technology for advanced combined cycle ramjet/scramjet engines being designed for the National Aerospace Plane (NASP). Technologies are being developed for both the dynamic seals that seal the sliding interfaces between articulating engine panels and sidewalls, and for the static seals that seal the heat exchanger to back-up structure interfaces. This viewgraph presentation provides an overview of the candidate engine seal concepts, seal material assessments, and unique test facilities used to assess the leakage and thermal performance of the seal concepts.

  18. Opposed Jet Burner Approach for Characterizing Flameholding Potentials of Hydrocarbon Scramjet Fuels

    NASA Technical Reports Server (NTRS)

    Pellett, Gerald L.; Convery, Janet L.; Wilson, Lloyd G.

    2006-01-01

    Opposed Jet Burner (OJB) tools have been used extensively by the authors to measure Flame Strength (FS) extinction limits of laminar H2/N2 air and (recently) hydrocarbon (HC) air Counterflow Diffusion Flames (CFDFs) at one atm. This paper details normalization of FSs of N2- diluted H2 and HC systems to account for effects of fuel composition, temperature, pressure, jet diameter, inflow Reynolds number, and inflow velocity profile (plug, contoured nozzle; and parabolic, straight tube). Normalized results exemplify a sensitive accurate means of validating, globally, reduced chemical kinetic models at approx. 1 atm and the relatively low temperatures approximating the loss of non-premixed idealized flameholding, e.g., in scramjet combustors. Laminar FS is defined locally as maximum air input velocity, U(sub air), that sustains combustion of a counter-jet of g-fuel at extinction. It uniquely characterizes a fuel. And global axial strain rate at extinction (U(sub air) normalized by nozzle or tube diameter, D(sub n or (sub t)) can be compared directly with computed extinction limits, determined using either a 1-D Navier Stokes stream-function solution, using detailed transport and finite rate chemistry, or (better yet) a detailed 2-D Navier Stokes numerical simulation. The experimental results define an idealized flameholding reactivity scale that shows wide ranging (50 x) normalized FS s for various vaporized-liquid and gaseous HCs, including, in ascending order: JP-10, methane, JP-7, n-heptane, n-butane, propane, ethane, and ethylene. Results from H2 air produce a unique and exceptionally strong flame that agree within approx. 1% of a recent 2-D numerically simulated FS for a 3 mm tube-OJB. Thus we suggest that experimental FS s and/or FS ratios, for various neat and blended HCs w/ and w/o additives, offer accurate global tests of chemical kinetic models at the Ts and Ps of extinction. In conclusion, we argue the FS approach is more direct and fundamental, for

  19. International Symposium on Air Breathing Engines, 10th, Nottingham, England, Sept. 1-6, 1991, Proceedings. Vols. 1 2

    SciTech Connect

    Billig, F.S.

    1991-01-01

    Among the topics discussed are high-speed transport, compressor aerodynamics, environment and pollution, engine performance, computational fluid dynamics, and combustion. Attention is also given to radial flow machines, hypersonic propulsion, engine condition monitoring, cascades and fans, inlets, tribology and materials, and transition and fluid dynamics. Also considered are mixing and mixing control; surge, stall, and flutter; combustion and aerothermodynamics; ram rockets, nozzles; icing and particles; nacelle design; supersonic combustion; scramjet and ramjet; turbines and heat transfer; and energy analysis.

  20. Numerical investigation of a hydrogen-fueled scramjet combustor at flight conditions

    NASA Astrophysics Data System (ADS)

    Rabadan, E.; Weigand, B.

    2013-03-01

    Numerical investigations of a hydrogen-fueled scramjet combustor at cruise flight conditions of Mach 8 at an altitude of 30 km have been performed. Two combustor configurations were investigated: a single-stage combustor with a central strut injector and a two-staged combustor combining the central strut and wall-mounted ramp injectors. These numerical simulations are aimed to study the flow structure, supersonic mixing, autoignition, and combustion for the present combustor configurations. A turbulent flow from a separate intake calculation was used as inflow condition for the combustor. A better performance for the two-staged combustor configuration was observed. The combination of the central strut injection together with the wall-ramp injection improved the turbulent mixing and, consequently, the combustion process. Inside the supersonic combustion chamber, the autoignition zone occurred downstream of the injectors, and combustion takes place accompanied by high heat release and pressure rise. As the equivalence ratio was increased, the combustion became stronger causing an upstream displacement of the shock train producing different pressure variations. For the two-staged combustor configuration, the location of the autoignition zone was found to appear further upstream compared to single-stage combustor. Mixing was improved by addition of the second-stage injection. The influence of the wall temperature was also investigated showing an effect on the combustion pressure rise and the length and location of the shock train.

  1. Nitric Oxide PLIF Visualization of Simulated Fuel-Air Mixing in a Dual-Mode Scramjet

    NASA Technical Reports Server (NTRS)

    Cantu, Luca M. L.; Gallo, Emanuela C. A.; Cutler, Andrew D.; Bathel, Brett F.; Danehy, Paul M.; Rockwell, Robert D.; Goyne, Christopher P.; McDaniel, James C.

    2015-01-01

    Nitric oxide (NO) planar induced laser fluorescence (PLIF) measurements have been performed in a small scale scramjet combustor at the University of Virginia Aerospace Research Laboratory at nominal simulated Mach 5 flight. A mixture of NO and N2 was injected at the upstream end of the inlet isolator as a surrogate for ethylene fuel, and the mixing of this fuel simulant was studied with and without a shock train. The shock train was produced by an air throttle, which simulated the blockage effects of combustion downstream of the cavity flame holder. NO PLIF signal was imaged in a plane orthogonal to the freestream at the leading edge of the cavity. Instantaneous planar images were recorded and analyzed to identify the most uniform cases, which were achieved by varying the location of the fuel injection and shock train. This method was used to screen different possible fueling configurations to provide optimized test conditions for follow-on combustion measurements using ethylene fuel. A theoretical study of the selected NO rotational transitions was performed to obtain a LIF signal that is linear with NO mole fraction and approximately independent of pressure and temperature.

  2. Computational parametric study of sidewall-compression scramjet inlet performance at Mach 10

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1993-01-01

    A computational parametric study of three-dimensional, sidewall-compression scramjet inlets was performed to identify the effects of geometric parameters on inlet performance. The parameters were the leading-edge sweep angle, varied between 30 and 60 deg, and the leading-edge position of the cowl, located at the throat and at two forward positions. A laminar boundary layer with cold-wall (T(sub wall) = 300 K (540 R)) boundary conditions was imposed. The parametric study was performed for a Mach number of 10 and a unit free-stream Reynolds number of 7.06 x 10(exp 6) per meter (2.15 x 10(exp 6) per foot) at a geometric contraction ratio of 5. The performance of each configuration was evaluated in terms of the mass capture, throat Mach number, total pressure recovery, kinetic energy efficiency, and internal compression. One computation of an unswept configuration was included as a baseline to determine the effects of introducing leading-edge sweep on the flow-field parameters. The purpose of the computational parametric study was to perform a trade-off of the effects of various parameters on the global performance of the inlet. Although no single optimal configuration emerged, trade-offs among the stated performance parameters identified a leading-edge sweep angle of 45 deg as possessing the most attractive performance characteristics.

  3. Mach 10 computational study of a three-dimensional scramjet inlet flow field

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1995-01-01

    The present work documents the computational results for a combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall-compression scramjet inlet configuration at Mach 10. The three-dimensional Navier-Stokes code SCRAMIN was chosen for the computational portion of the study because it uses a well-known and well-proven numerical scheme and has shown favorable comparison with experiment at Mach numbers between 2 and 6. One advantage of CFD was that it provided flow field data for a detailed examination of the internal flow characteristics in addition to the surface properties. The experimental test matrix at mach 10 included three geometric contraction ratios (3, 5, and 9), three Reynolds numbers (0.55 x 10(exp 6) per foot, 1.14 x 10(exp 6) per foot, and 2.15 x 10(exp 6) per foot), and three cowl positions (at the throat and two forward positions). Computational data for two of these configurations (the contraction ratio of 3, Re = 2.15 x 10 (exp 6) per foot, at two cowl positions) are presented along with a detailed analysis of the flow interactions in successive computational planes.

  4. Mach 10 computational study of a three-dimensional scramjet inlet flow field

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1995-01-01

    The present work documents the computational results for a combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall-compression scramjet inlet configuration at Mach 10. The three-dimensional Navier-Stokes code SCRAMIN was chosen for the computational portion of the study because it uses a well-known and well-proven numerical scheme and has shown favorable comparison with experiment at Mach numbers between 2 and 6. One advantage of CFD was that it provided flow field data for a detailed examination of the internal flow characteristics in addition to the surface properties. The experimental test matrix at Mach 10 included three geometric contraction ratios (3, 5, and 9), three Reynolds numbers (0.55 x 10(exp 6) per foot, 1.14 x 10(exp 6) per foot, and 2.15 x 10(exp 6) per foot), and three cowl positions (at the throat and two forward positions). Computational data for two of these configurations (the contraction ratio of 3, Re = 2.15 x 10(exp 6) per foot, at two cowl positions) are presented along with a detailed analysis of the flow interactions in successive computational planes.

  5. Characterization of a swept-strut hydrogen fuel-injector for scramjet applications

    NASA Technical Reports Server (NTRS)

    Northam, G. B.; Trexler, C. A.; Anderson, G. Y.

    1978-01-01

    Results of an experimental investigation of a swept-strut hydrogen fuel-injector simulating the center strut of a three strut scramjet module at Mach 6 flight conditions are presented. Detailed wall pressure distributions from over 100 separate tests with overall fuel flow from 0.1 to 1.3 times stoichiometric and test gas stagnation temperature from 1100 to 2400 K were recorded. The distance for pressure rise from the point of injection was found to increase with increasing test stagnation temperature. This trend indicates that chemical kinetics in the immediate region of perpendicular injection are not likely to be the mechanism controlling the onset of pressure rise. A fluid dynamic mechanism is suggested involving separation of the boundary layer downstream of injection which is forced upstream from the trailing edge by pressure rise due to combustion occurring in the base region of the strut. The results obtained indicate that the swept-strut fuel-injector concept can be adapted to a wide range of flight conditions by varying the amount of perpendicular fuel injection.

  6. Hydrogen-fueled scramjet investigation. Final report, Feb 87-Apr 90

    SciTech Connect

    Peschke, W.T.; Barber, T.J.; Chiappetta, L.; Anderson, T.J.; Eckerle, W.J.

    1991-02-01

    This report documents an experimental and analytical study of the supersonic mixing and combustion processes associated with a free-shear layer. A specific objective comprised quantifying the effects on the mixing and combustion processes of pressure gradients associated with the production of shock waves in the flowfield. An experiment was conducted in a generic supersonic combustion configuration that incorporated the essential characteristics of scramjet combustors currently being considered for flight application. The objective of the analytical effort was to evaluate mixing and combustion models in state-of-the-art computational fluid dynamics (CFD) codes. The experiments provided accurate data used to formulate code input conditions and to evaluate codes that were formulated for application in current high-speed propulsion design activities. Species concentration, total and static temperature, and pitot pressure profiles were measured in a shear layer experiment in which gaseous hydrogen was injected tangentially into a Mach 3 (nominal) vitiated air stream. In this experiment, hydrogen was injected adjacent to the upperwall of the test section. A nonintrusive optical diagnostic technique was used in conjunction with intrusive probing to define the flowfield initial conditions and to characterize the detailed features of the flowfield subsequent to hydrogen injection and combustion.

  7. Copper contamination effects on hydrogen-air combustion under SCRAMJET (supersonic combustion ramjet) testing conditions

    SciTech Connect

    Chang, S.L.; Lottes, S.A.; Berry, G.F.

    1990-01-01

    Two forms of copper catalytic reactions (homogeneous and heterogeneous) in hydrogen flames were found in a literature survey. Hydrogen atoms in flames recombine into hydrogen molecules through catalytic reactions, and these reactions which affect the timing of the combustion process. Simulations of hydrogen flames with copper contamination were conducted by using a modified general chemical kinetics program (GCKP). Results show that reaction times of hydrogen flames are shortened by copper catalytic reactions, but ignition times are relatively insensitive to the reactions. The reduction of reaction time depends on the copper concentration, copper phase, particle size (if copper is in the condensed phase), and initial temperature and pressure. The higher the copper concentration of the smaller the particle, the larger the reduction in reaction time. For a supersonic hydrogen flame (Mach number = 4.4) contaminated with 200 ppm of gaseous copper species, the calculated reaction times are reduced by about 9%. Similar reductions in reaction time are also computed for heterogeneous copper contamination. Under scramjet testing conditions, the change of combustion timing appears to be tolerable (less than 5%) if the Mach number is lower than 3 or the copper contamination is less than 100 ppm. The higher rate the Mach number, the longer the reaction time and the larger the copper catalytic effects. 7 tabs., 8 figs., 34 refs.

  8. Hysteresis of mode transition in a dual-struts based scramjet

    NASA Astrophysics Data System (ADS)

    Yan, Zhang; Shaohua, Zhu; Bing, Chen; Xu, Xu

    2016-11-01

    Tests and numerical simulations were performed to investigate the combustion performance of a dual-staged scramjet combustor. High enthalpy vitiated inflow at a total temperature of 1231 K was supplied using a hydrogen-combustion heater. The inlet Mach number was 2.0. Liquid kerosene was injected into the combustor using the dual crossed struts. Three-dimensional Reynolds averaged reacting flow was solved using a two-equation k-ω SST turbulence model to calculate the effect of turbulent stress, and a partial-premixed flamelet model to model the effects of turbulence-chemistry interactions. The discrete phase model was utilized to simulate the fuel atomization and vaporization. For simplicity, the n-decane was chosen as the surrogate fuel with a reaction mechanism of 40 species and 141 steps. The predicted wall pressure profiles at three fuel injection schemes basically captured the axial varying trend of the experimental data. With the downstream equivalence ratio held constant, the upstream equivalence ratio was numerically increased from 0.1 to 0.4 until a steady combustion was obtained. Subsequently, the upstream equivalence ratio was decreased from 0.4 to 0.1 once again. Two ramjet modes with different wall pressure profiles and corresponding flow structures were captured under the identical upstream equivalence ratio of 0.1, illustrating an obvious hysteresis phenomenon. The mechanism of this hysteresis was explained by the transition hysteresis of the pre-combustion shock train in the isolator.

  9. OH PLIF Visualization of a Premixed Ethylene-fueled Dual-Mode Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Cantu, Luca M. L.; Gallo, Emanuela C. A.; Cutler, Andrew D.; Danehy, Paul M.; Johansen, Craig T.; Rockwell, Robert D.; Goyne, Christopher P.; McDaniel, James C.

    2016-01-01

    Hydroxyl radical (OH) planar induced laser fluorescence (PLIF) measurements have been performed in a small-scale scramjet combustor at the University of Virginia Aerospace Research Laboratory at nominal simulated Mach 5 enthalpy. OH lines were carefully chosen to have fluorescent signal that is independent of pressure and temperature but linear with mole fraction. The OH PLIF signal was imaged in planes orthogonal to and parallel to the freestream flow at different equivalence ratios. Flameout limits were tested and identified. Instantaneous planar images were recorded and analyzed to compare the results with width increased dual-pump enhanced coherent anti-Stokes Raman spectroscopy (WIDECARS) measurements in the same facility and large eddy simulation/Reynolds average Navier-Stokes (LES/RANS) numerical simulation. The flame angle was found to be approximately 10 degrees for several different conditions, which is in agreement with numerical predictions and measurements using WIDECARS. Finally, a comparison between NO PLIF non-combustion cases and OH PLIF combustion cases is provided: the comparison reveals that the dominant effect of flame propagation is freestream turbulence rather than heat release and concentration gradients.

  10. 1-D Numerical Analysis of RBCC Engine Performance

    NASA Technical Reports Server (NTRS)

    Han, Samuel S.

    1998-01-01

    An RBCC engine combines air breathing and rocket engines into a single engine to increase the specific impulse over an entire flight trajectory. Considerable research pertaining to RBCC propulsion was performed during the 1960's and these engines were revisited recently as a candidate propulsion system for either a single-stage-to-orbit (SSTO) or two-stage-to-orbit (TSTO) launch vehicle. There are a variety of RBCC configurations that had been evaluated and new designs are currently under development. However, the basic configuration of all RBCC systems is built around the ejector scramjet engine originally developed for the hypersonic airplane. In this configuration, a rocket engine plays as an ejector in the air-augmented initial acceleration mode, as a fuel injector in scramjet mode and the rocket in all rocket mode for orbital insertion. Computational fluid dynamics (CFD) is a useful tool for the analysis of complex transport processes in various components in RBCC propulsion systems. The objective of the present research was to develop a transient 1-D numerical model that could be used to predict flow behavior throughout a generic RBCC engine following a flight path.

  11. Quantitative measurement of transverse injector and free stream interaction in a nonreacting SCRAMJET combustor using laser-induced iodine fluorescence

    NASA Technical Reports Server (NTRS)

    Fletcher, D. G.; Mcdaniel, J. C.

    1987-01-01

    A preliminary quantitative study of the compressible flowfield in a steady, nonreacting model SCRAMJET combustor using laser-induced iodine fluorescence (LIIF) is reported. Measurements of density, temperature, and velocity were conducted with the calibrated, nonintrusive, optical technique for two different combustor operating conditions. First, measurements were made in the supersonic flow over a rearward-facing step without transverse injection for comparison with calculated pressure profiles. The second configuration was staged injection behind the rearward-facing step at an injection dynamic pressure ratio of 1.06. These experimental results will be used to validate computational fluid dynamic (CFD) codes being developed to model supersonic combustor flowfields.

  12. Hyper-X Research Vehicle - Artist Concept Mounted on Pegasus Rocket Attached to B-52 Launch Aircraft

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This artist's concept depicts the Hyper-X research vehicle riding on a booster rocket prior to being launched by the Dryden Flight Research Center's B-52 at about 40,000 feet. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry

  13. Pegasus Rocket Booster Being Prepared for X-43A/Hyper-X Flight Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

    Technicians prepare a Pegasus rocket booster for flight tests with the X-43A 'Hypersonic Experimental Vehicle,' or 'Hyper-X.' The X-43A, which will be attached to the Pegasus booster and drop launched from NASA's B-52 mothership, was developed to research dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet

  14. Pegasus Rocket Booster Being Prepared for X-43A/Hyper-X Flight Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

    A close-up view of the front end of a Pegasus rocket booster being prepared by technicians at the Dryden Flight Research Center for flight tests with the X-43A 'Hypersonic Experimental Vehicle,' or 'Hyper-X.' The X-43A, which will be attached to the Pegasus booster and drop launched from NASA's B-52 mothership, was developed to research dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are

  15. Hyper-X Vehicle Model - Top Rear View

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This aft-quarter model view of NASA's X-43A 'Hyper-X' or Hypersonic Experimental Vehicle shows its sleek, geometric design. The X-43A was developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen

  16. X-43A/Hyper-X Vehicle Arrives at NASA Dryden

    NASA Technical Reports Server (NTRS)

    1999-01-01

    A close-up of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X,' in its protective shipping framework as it arrives at the Dryden Flight Research Center in October 1999. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only

  17. X-43A Vehicle During Ground Testing

    NASA Technical Reports Server (NTRS)

    1999-01-01

    The X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' is seen here undergoing ground testing at NASA's Dryden Flight Research Center, Edwards, California in December 1999. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only

  18. X-43A Vehicle During Ground Testing

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This photo shows a close-up, rear view of the X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' undergoing ground testing at NASA's Dryden Flight Research Center, Edwards, California in December 1999. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology

  19. X-43A Vehicle During Ground Testing

    NASA Technical Reports Server (NTRS)

    1999-01-01

    The X-43A Hypersonic Experimental Vehicle, or 'Hyper-X' is seen here undergoing ground testing at NASA's Dryden Flight Research Center, Edwards, California. The X-43A was developed to research a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By

  20. Hyper-X Vehicle Model - Front View

    NASA Technical Reports Server (NTRS)

    1996-01-01

    A front view of an early desk-top model of NASA's X-43A 'Hyper-X,' or Hypersonic Experimental Vehicle, which has been developed to flight test a dual-mode ramjet/scramjet propulsion system at speeds from Mach 7 up to Mach 10 (7 to 10 times the speed of sound, which varies with temperature and altitude). Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic

  1. Ramjet engine test facility (RJTF). Technical report

    SciTech Connect

    1998-02-01

    The National Aerospace Laboratory of Japan constructed a ramjet engine test facility (RJTF) at the Kakuda Research Center in 1994. It can duplicate engine test conditions in the range of flight Mach numbers from 4 to 8. The facility can supply non-vitiated air for M4 and M6 to identify the contamination effect in the vitiated air, to provide the basis for evaluating engine performance in the M8 flight condition. This paper outlines the unique features and operating characteristics of the RJTF. The quality of air stream obtained during facility calibration, and the facility-engine interaction are described. Finally the authors review tests of an H2-fueled scramjet that are currently underway.

  2. Two-Dimensional Scramjet Inlet Unstart Model: Wind-Tunnel Blockage and Actuation Systems Test

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1994-01-01

    This supplement to NASA TM 109152 shows the Schlieren video (10 min. 52 sec., color, Beta and VHS) of the external flow field and a portion of the internal flow field of a two-dimensional scramjet inlet model in the NASA Langley 20-Inch Mach 6 Tunnel. The intent of the overall test program is to study (both experimentally and computationally) the dynamics of the inlet unstart; this (phase I) effort examines potential wind-tunnel blockage issues related to model sizing and the adequacy of the actuation systems in accomplishing the start and unstart. The model is equipped with both a moveable cowl and aft plug. Windows in the inlet sidewalls allow limited optical access to the internal shock structure. In the video, flow is from right to left, and the inlet is oriented inverted with respect to flight, i.e., with the cowl on top. The plug motion is obvious because the plug is visible in the aft window. The cowl motion, however, is not as obvious because the cowl is hidden from view by the inlet sidewall. The end of the cowl actuator arm, however, becomes visible above the inlet sidewalls between the windows when the cowl is up (see figure 1b of the primary document). The model is injected into the tunnel and observed though several actuation sequences with two plug configurations over a range of unit freestream Reynolds number at a nominal freestream Mach number of 6. The framing rate and shutter speed of the camera were too slow to fully capture the dynamics of the unstart but did prove sufficient to identify inlet start and unstart. This series of tests indicated that the model was appropriately sized for this facility and identified operability limits required first to allow the inlet to start and second to force the unstart.

  3. Effects of nonuniform Mach-number entrance on scramjet nozzle flowfield and performance

    NASA Astrophysics Data System (ADS)

    Zhang, Pu; Xu, Jinglei; Quan, Zhibin; Mo, Jianwei

    2016-12-01

    Considering the non-uniformities of nozzle entrance influenced by the upstream, the effects of nonuniform Mach-number coupled with shock and expansion-wave on the flowfield and performances of single expansion ramp nozzle (SERN) are numerically studied using Reynolds-Averaged Navier-Stokes equations. The adopted Reynolds-averaged Navier-Stokes methodology is validated by comparing the numerical results with the cold experimental data, and the average method used in this paper is discussed. Uniform and nonuniform facility nozzles are designed to generate different Mach-number profile for the inlet of SERN, which is direct-connected with different facility nozzle, and the whole flowfield is simulated. Because of the coupling of shock and expansion-wave, flow direction of nonuniform SERN entrance is distorted. Compared with Mach contour of uniform case, the line is more curved for coupling shock-wave entrance (SWE) case, and flatter for the coupling expansion-wave entrance (EWE) case. Wall pressure distribution of SWE case appears rising region, whereas decreases like stairs of EWE case. The numerical results reveal that the coupled shock and expansion-wave play significant roles on nozzle performances. Compared with the SERN performances of uniform entrance case at the same work conditions, the thrust of nonuniform entrance cases reduces by 3-6%, pitch moment decreases by 2.5-7%. The negative lift presents an incremental trend with EWE while the situation is the opposite with SWE. These results confirm that considering the entrance flow parameter nonuniformities of a scramjet nozzle coupled with shock or expansion-wave from the upstream is necessary.

  4. Characteristics of transverse hydrogen jet in presence of multi air jets within scramjet combustor

    NASA Astrophysics Data System (ADS)

    Barzegar Gerdroodbary, M.; Fallah, Keivan; Pourmirzaagha, H.

    2017-03-01

    In this article, three-dimensional simulation is performed to investigate the effects of micro air jets on mixing performances of cascaded hydrogen jets within a scramjet combustor. In order to compare the efficiency of this technique, constant total fuel rate is injected through one, four, eight and sixteen arrays of portholes in a Mach 4.0 crossflow with a fuel global equivalence ratio of 0.5. In this method, micro air jets are released within fuel portholes to augment the penetration in upward direction. Extensive studies were performed by using the Reynolds-averaged Navier-Stokes equations with Menter's Shear Stress Transport (SST) turbulence model. Numerical studies on various air and fuel arrangements are done and the mixing rate and penetration are comprehensively investigated. Also, the flow feature of the fuel and air jets for different configuration is revealed. According to the obtained results, the influence of the micro air jets is significant and the presence of micro air jets increases the mixing rate about 116%, 77%, 56% and 41% for single, 4, 8 and 16 multi fuel jets, respectively. The maximum mixing rate of the hydrogen jet is obtained when the air jets are injected within the sixteen multi fuel jets. According to the circulation analysis of the flow for different air and fuel arrangements, it was found that the effects of air jets on flow structure are varied in various conditions and the presence of the micro jet highly intensifies the circulation in the case of 8 and 16 multi fuel jets.

  5. Mach 10 experimental database of a three-dimensional scramjet inlet flow field

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1995-01-01

    The present work documents the experimental database of a combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall compression scramjet inlet configuration at Mach 10. A total of 356 channels of pressure data, including static pressure orifices, pitot pressures, and exit flow rakes, along with oil flow and infrared thermography, provided a detailed experimental description of the flow. Mach 10 tests were performed for three geometric contraction ratios (3, 5, and 9), three Reynolds numbers (0.55 x 10(exp 6) per foot, 1.14 x 10(exp 6) per foot, and 2.15 x 10(exp 6) per foot), and three cowl positions (at the throat and two forward positions). For the higher contraction ratios, a large forward separation of the inflow boundary layer was observed, making the high contraction ratio configurations unsuitable for flight operation. A decrease in the freestream unit Reynolds number (Re) of only a factor of 2 led to a similar upstream separation. Although the presence of such large-scale separations leads to the question of whether the inlet is started, the presence of internal oblique swept shock interactions on the sidewalls seems to indicate that at least in the classical sense, the inlet is not unstarted. The laminar inflow boundary layer therefore appears to be very sensitive to increases in contraction ratio (CR) or decreases in Reynolds number; only the CR = 3 configuration with 0.25, and 50 percent cowl at Re = 2.15 x 10(exp 6) per foot operated 'on design'.

  6. A Priori Analysis of a Compressible Flamelet Model using RANS Data for a Dual-Mode Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Quinlan, Jesse R.; Drozda, Tomasz G.; McDaniel, James C.; Lacaze, Guilhem; Oefelein, Joseph

    2015-01-01

    In an effort to make large eddy simulation of hydrocarbon-fueled scramjet combustors more computationally accessible using realistic chemical reaction mechanisms, a compressible flamelet/progress variable (FPV) model was proposed that extends current FPV model formulations to high-speed, compressible flows. Development of this model relied on observations garnered from an a priori analysis of the Reynolds-Averaged Navier-Stokes (RANS) data obtained for the Hypersonic International Flight Research and Experimentation (HI-FiRE) dual-mode scramjet combustor. The RANS data were obtained using a reduced chemical mechanism for the combustion of a JP-7 surrogate and were validated using avail- able experimental data. These RANS data were then post-processed to obtain, in an a priori fashion, the scalar fields corresponding to an FPV-based modeling approach. In the current work, in addition to the proposed compressible flamelet model, a standard incompressible FPV model was also considered. Several candidate progress variables were investigated for their ability to recover static temperature and major and minor product species. The effects of pressure and temperature on the tabulated progress variable source term were characterized, and model coupling terms embedded in the Reynolds- averaged Navier-Stokes equations were studied. Finally, results for the novel compressible flamelet/progress variable model were presented to demonstrate the improvement attained by modeling the effects of pressure and flamelet boundary conditions on the combustion.

  7. Fabrication and development of several heat pipe honeycomb sandwich panel concepts

    NASA Astrophysics Data System (ADS)

    Tanzer, H. J.

    1982-06-01

    The feasibility of fabricating and processing liquid metal heat pipes in a low mass honeycomb sandwich panel configuration for application on the NASA Langley airframe-integrated Scramjet engine was investigated. A variety of honeycomb panel facesheet and core-ribbon wick concepts was evaluated within constraints dictated by existing manufacturing technology and equipment. The chosen design consists of an all-stainless steel structure, sintered screen facesheets, and two types of core-ribbon; a diffusion bonded wire mesh and a foil-screen composite. Cleaning, fluid charging, processing, and process port sealing techniques were established. The liquid metals potassium, sodium and cesium were used as working fluids. Eleven honeycomb panels 15.24 cm X 15.24 cm X 2.94 cm were delivered to NASA Langley for extensive performance testing and evaluation; nine panels were processed as heat pipes, and two panels were left unprocessed.

  8. Test description and preliminary pitot-pressure surveys for Langley Test Technique Demonstrator at Mach 6

    NASA Technical Reports Server (NTRS)

    Everhart, Joel L.; Ashby, George C., Jr.; Monta, William J.

    1992-01-01

    A propulsion/airframe integration experiment conducted in the NASA Langley 20-Inch Mach 6 Tunnel using a 16.8-in.-long version of the Langley Test Technique Demonstrator configuration with simulated scramjet propulsion is described. Schlieren and vapor screen visualization of the nozzle flow field is presented and correlated with pitot-pressure flow-field surveys. The data were obtained at nominal free-stream conditions of Re = 2.8 x 10 exp 6 and a nominal engine total pressure of 100 psia. It is concluded that pitot-pressure surveys coupled to schlieren and vapor-screen photographs, and oil flows have revealed flow features including vortices, free shear layers, and shock waves occurring in the model flow field.

  9. Design, fabrication and test of liquid metal heat-pipe sandwich panels

    NASA Technical Reports Server (NTRS)

    Basiulis, A.; Camarda, C. J.

    1983-01-01

    Integral heat-pipe sandwich panels, which synergistically combine the thermal efficiency of heat pipes and the structural efficiency of honeycomb sandwich panel construction, were fabricated and tested. The designs utilize two different wickable honeycomb cores, facesheets with screen mesh sintered to the internal surfaces, and potassium or sodium as the working fluid. Panels were tested by radiant heating, and the results indicate successful heat pipe operation at temperatures of approximately 922K (1200F). These panels, in addition to solving potential thermal stress problems in an Airframe-Integrated Scramjet Engine, have potential applications as cold plates for electronic component cooling, as radiators for space platforms, and as low distortion, large area structures.

  10. Visualization and Analysis of a Hydrocarbon Premixed Flame a in Small Scale Scramjet Combustor

    NASA Astrophysics Data System (ADS)

    Cantu, Luca Maria Luigi

    Nitric oxide (NO) planar induced laser fluorescence (PLIF) measurements have been performed in a small scale scramjet combustor at the University of Virginia Aerospace Research Laboratory at nominal simulated Mach 5 flight enthalpy. A mixture of NO and N2 was injected at the upstream end of the inlet isolator as a surrogate for ethylene fuel, and the mixing of this fuel simulant was studied with and without a shock train. The shock train was produced by an air throttle, which simulated the blockage effects of combustion downstream of the cavity flame holder. NO PLIF signal was imaged in a plane orthogonal to the freestream at the leading edge of the cavity. Instantaneous planar images were recorded and analyzed to identify the most uniform cases, which were achieved by varying the location of the fuel injection and shock train. This method was used to screen different possible fueling configurations to provide optimized test conditions for follow-on combustion measurements using ethylene fuel. A theoretical study of the selected NO rotational transitions was performed to obtain a LIF signal that is linear with NO mole fraction and approximately independent of pressure and temperature. In the same facility, OH PLIF measurements were also performed; OH lines were carefully chosen to have fluorescent signal that is independent of pressure and temperature but linear with mole fraction. The OH PLIF signal was imaged in planes orthogonal to and parallel to the freestream flow at different equivalence ratios. Flameout limits were tested and identified. Instantaneous planar images were recorded and analyzed to compare the results with width increased dual-pump enhanced coherent anti-Stokes Raman spectroscopy (WIDECARS) measurements in the same facility and large eddy simulation/Reynolds average Navier-Stokes (LES/RANS) numerical simulations. The flame angle was found to be approximately 10 degrees for several different conditions, which is in agreement with numerical

  11. Characterization of the NASA Langley Arc Heated Scramjet Test Facility Using NO PLIF

    NASA Technical Reports Server (NTRS)

    Kidd, F. Gray, III; Narayanaswamy, Venkateswaran; Danehy, Paul M.; Inman, Jennifer A.; Bathel, Brett F.; Cabell, Karen F.; Hass, Neal E.; Capriotti, Diego P.; Drozda, Tomasz G.; Johansen, Criag T.

    2014-01-01

    The nitric oxide planar laser-induced fluorescence (NO PLIF) imaging was used to characterize the air flow of the NASA Langley Arc Heated Scramjet Test Facility (AHSTF) configured with a Mach 6 nozzle. The arc raises the enthalpy of the test gas in AHSTF, producing nitric oxide. Nitric oxide persists as the temperature drops through the nozzle into the test section. NO PLIF was used to qualitatively visualize the flowfield at different experimental conditions, measure the temperature of the gas flow exiting the facility nozzle, and visualize the wave structure downstream of the nozzle at different operating conditions. Uniformity and repeatability of the nozzle flow were assessed. Expansion and compression waves on the free-jet shear layer as the nozzle flow expands into the test section were visualized. The main purpose of these experiments was to assess the uniformity of the NO in the freestream gas for planned experiments, in which NO PLIF will be used for qualitative fuel-mole-fraction sensitive imaging. The shot-to-shot fluctuations in the PLIF signal, caused by variations in the overall laser intensity as well as NO concentration and temperature variations in the flow was 20-25% of the mean signal, as determined by taking the standard deviation of a set of images obtained at constant conditions and dividing by the mean. The fluctuations within individual images, caused by laser sheet spatial variations as well as NO concentration and temperature variations in the flow, were about 28% of the mean in images, determined by taking standard deviation within individual images, dividing by the mean in the same image and averaged over the set of images. Applying an averaged laser sheet intensity correction reduced the within-image intensity fluctuations to about 10% suggesting that the NO concentration is uniform to within 10%. There was no significant difference in flow uniformity between the low and high enthalpy settings. While not strictly quantitative, the

  12. Influencing factors on the mode transition in a dual-mode scramjet

    NASA Astrophysics Data System (ADS)

    Yan, Zhang; Bing, Chen; Gang, Liu; Baoxi, Wei; Xu, Xu

    2014-10-01

    An experimental investigation was performed to characterize the effects of fuel type, injector configuration, inflow total temperature and fuel injection distribution on the mode transition in a dual-mode scramjet combustor. High enthalpy vitiated air was heated to three total temperatures by the hydrogen-oxygen combustion, entering the isolator entrance at a Mach number of 2.0. Fuel was injected through a four-hole aero-ramp or transverse injector, and ignited by a gas-pilot flame. At an inflow stagnation condition of P0=0.85 MPa and T0=1200 K, three combustion modes, namely Pure Scram Mode, Dual-Mode Scram Mode and Dual-Mode Ram Mode, were classified through the wall pressure distributions, one-dimensional performance analysis, and optical visualization. Two critical fuel equivalence ratios were selected to divide three combustion modes. At the lower transition point, ERlc, the transition from Pure Scram Mode to Dual-Mode Scram Mode occurred corresponding to a normalized wall pressure at x/H=2.5 of 0.23; while at the upper transition point, ERuc, the transition from Dual-Mode Scram Mode to Dual-Mode Ram Mode occurred corresponding to a normalized wall pressure at x/H=2.5 of 0.34. The transition width, ERw, was defined as the difference of two transition points. In this limited range, the combustor was operating in the Dual-Mode Scram Mode. The ERuc was estimated based on the Rayleigh flow relation, and the applicability of the analytic equation was testified through a series of experiments with different boundary conditions. The experimental results showed that two transition points in the ethylene case were higher in ER than in the hydrogen case, and the measured ERuc of two fuels were all 2.5 times larger than the predicted values. However, two fuels had little difference on the transition width. Due to the angled injection, two transition points in the aero-ramp case were higher in ER than in the transverse injector case; while the influence of injector

  13. Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

    NASA Technical Reports Server (NTRS)

    Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

    2002-01-01

    This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

  14. Status of the Combined-Cycle Engine Large-Scale Inlet Mode Transition Experiment

    NASA Technical Reports Server (NTRS)

    Slater, John W.; Saunders, J. Dave

    2011-01-01

    The Large-Scale Inlet Mode Transition (LIMX) experiment is currently being conducted in the 10x10 foot supersonic wind tunnel at the NASA Glenn Research Center. The experiment has involved the efforts of a team for over four years to get to the first phase of testing, which is examining the aerodynamic characteristics of the inlet. The LIMX inlet involves dual flowpaths: one to provide flow to a turbine engine and one to provide flow to a dual-mode ramjet/scramjet. A rotating splitter cowl can close off the turbine flowpath, which would occur as the propulsion system transitions from turbine power to ramjet/scramjet power at Mach 4. The first phase of the experiment will simulate the turbine and ramjet/scramjet flows using cold pipes with flow rates controlled by mass-flow plugs. Much of the testing will characterize the performance of the turbine flowpath (total pressure recovery and distortion at the engine face location) as factors such as bleed rates and configuration and vortex generators are varied during the inlet mode transition. The performance of the inlet will also be examined at off-design Mach numbers (2.5-3.0) and at angle-of-attack.

  15. Hyper-X and Pegasus Launch Vehicle: A Three-Foot Model of the Hypersonic Experimental Research Vehic

    NASA Technical Reports Server (NTRS)

    1997-01-01

    A close-up view of the X-43A Hypersonic Experimental Research Vehicle, or Hyper-X, portion of a three-foot-long model of the vehicle/booster combination at NASA's Dryden Flight Research Center, Edwards, California. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry heavier payloads. Another unique aspect of the X-43A vehicle is

  16. Hyper-X and Pegasus Launch Vehicle: A Three-Foot Model of the Hypersonic Experimental Research Vehic

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The configuration of the X-43A Hypersonic Experimental Research Vehicle, or Hyper-X, attached to a Pegasus launch vehicle is displayed in this side view of a three-foot-long model of the vehicle/booster combination at NASA's Dryden Flight Research Center, Edwards, California. Hyper-X, the flight vehicle for which is designated as X-43A, is an experimental flight-research program seeking to demonstrate airframe-integrated, 'air-breathing' engine technologies that promise to increase payload capacity for future vehicles, including hypersonic aircraft (faster than Mach 5) and reusable space launchers. This multiyear program is currently underway at NASA Dryden Flight Research Center, Edwards, California. The Hyper-X schedule calls for its first flight later this year (2000). Hyper-X is a joint program, with Dryden sharing responsibility with NASA's Langley Research Center, Hampton, Virginia. Dryden's primary role is to fly three unpiloted X-43A research vehicles to validate engine technologies and hypersonic design tools as well as the hypersonic test facility at Langley. Langley manages the program and leads the technology development effort. The Hyper-X Program seeks to significantly expand the speed boundaries of air-breathing propulsion by being the first aircraft to demonstrate an airframe-integrated, scramjet-powered free flight. Scramjets (supersonic-combustion ramjets) are ramjet engines in which the airflow through the whole engine remains supersonic. Scramjet technology is challenging because only limited testing can be performed in ground facilities. Long duration, full-scale testing requires flight research. Scramjet engines are air-breathing, capturing their oxygen from the atmosphere. Current spacecraft, such as the Space Shuttle, are rocket powered, so they must carry both fuel and oxygen for propulsion. Scramjet technology-based vehicles need to carry only fuel. By eliminating the need to carry oxygen, future hypersonic vehicles will be able to carry

  17. Gaseous Surrogate Hydrocarbons for a Hifire Scramjet that Mimic Opposed Jet Extinction Limits for Cracked JP Fuels

    NASA Technical Reports Server (NTRS)

    Pellett, Gerald L.; Vaden, Sarah N.; Wilson, Lloyd G.

    2008-01-01

    This paper describes, first, the top-down methodology used to define simple gaseous surrogate hydrocarbon (HC) fuel mixtures for a hypersonic scramjet combustion subtask of the HiFIRE program. It then presents new and updated Opposed Jet Burner (OJB) extinction-limit Flame Strength (FS) data obtained from laminar non-premixed HC vs. air counterflow diffusion flames at 1-atm, which follow from earlier investigations. FS represents a strain-induced extinction limit based on cross-section-average air jet velocity, U(sub air), that sustains combustion of a counter jet of gaseous fuel just before extinction. FS uniquely characterizes a kinetically limited fuel combustion rate. More generally, Applied Stress Rates (ASRs) at extinction (U(sub air) normalized by nozzle or tube diameter, D(sub n or t) can directly be compared with extinction limits determined numerically using either a 1-D or (preferably) a 2-D Navier Stokes simulation with detailed transport and finite rate chemistry. The FS results help to characterize and define three candidate surrogate HC fuel mixtures that exhibit a common FS 70% greater than for vaporized JP-7 fuel. These include a binary fuel mixture of 64% ethylene + 36% methane, which is our primary recommendation. It is intended to mimic the critical flameholding limit of a thermally- or catalytically-cracked JP-7 like fuel in HiFIRE scramjet combustion tests. Our supporting experimental results include: (1) An idealized kinetically-limited ASR reactivity scale, which represents maximum strength non-premixed flames for several gaseous and vaporized liquid HCs; (2) FS characterizations of Colket and Spadaccini s suggested ternary surrogate, of 60% ethylene + 30% methane + 10% n-heptane, which matches the ignition delay of a typical cracked JP fuel; (3) Data showing how our recommended binary surrogate, of 64% ethylene + 36% methane, has an identical FS; (4) Data that characterize an alternate surrogate of 44% ethylene + 56% ethane with identical

  18. Effect of cavity configuration on kerosene spark ignition in a scramjet combustor at Ma 4.5 flight condition

    NASA Astrophysics Data System (ADS)

    Bao, Heng; Zhou, Jin; Pan, Yu

    2015-12-01

    Spark ignition experiments of liquid kerosene are conducted in a scramjet model equipped with dual-cavities at Ma 4.5 flight condition with a stagnation temperature of 1032 K. The ignition ability of two cavities with different length is compared and analyzed based on the wall pressure distribution along the combustor and the thrust evolution. The experimental results indicate that the longer cavity (L/D=7) is more suitable than the smaller cavity (L/D=5) in spark ignition. When employing the smaller cavity, three steady combustion states are observed after spark ignition. The concept of 'local flame' is adopted to explain the expanding problem of weak combustion. The local equivalence ratio in the shear layer is the dominated factor in determining the developing process of local flame. The final steady combustion mode of the combustor is dependent on the flame developing process. When employing the longer cavity, the establishment of intense combustion state can be much easier.

  19. Developing Conceptual Hypersonic Airbreathing Engines Using Design of Experiments Methods

    NASA Technical Reports Server (NTRS)

    Ferlemann, Shelly M.; Robinson, Jeffrey S.; Martin, John G.; Leonard, Charles P.; Taylor, Lawrence W.; Kamhawi, Hilmi

    2000-01-01

    Designing a hypersonic vehicle is a complicated process due to the multi-disciplinary synergy that is required. The greatest challenge involves propulsion-airframe integration. In the past, a two-dimensional flowpath was generated based on the engine performance required for a proposed mission. A three-dimensional CAD geometry was produced from the two-dimensional flowpath for aerodynamic analysis, structural design, and packaging. The aerodynamics, engine performance, and mass properties arc inputs to the vehicle performance tool to determine if the mission goals were met. If the mission goals were not met, then a flowpath and vehicle redesign would begin. This design process might have to be performed several times to produce a "closed" vehicle. This paper will describe an attempt to design a hypersonic cruise vehicle propulsion flowpath using a Design of' Experiments method to reduce the resources necessary to produce a conceptual design with fewer iterations of the design cycle. These methods also allow for more flexible mission analysis and incorporation of additional design constraints at any point. A design system was developed using an object-based software package that would quickly generate each flowpath in the study given the values of the geometric independent variables. These flowpath geometries were put into a hypersonic propulsion code and the engine performance was generated. The propulsion results were loaded into statistical software to produce regression equations that were combined with an aerodynamic database to optimize the flowpath at the vehicle performance level. For this example, the design process was executed twice. The first pass was a cursory look at the independent variables selected to determine which variables are the most important and to test all of the inputs to the optimization process. The second cycle is a more in-depth study with more cases and higher order equations representing the design space.

  20. Subscale Testing of a Ceramic Composite Cooled Panel Led to Its Design and Fabrication for Scramjet Engine Testing

    NASA Technical Reports Server (NTRS)

    Jaskowiak, Martha H.

    2004-01-01

    In a partnership between the NASA Glenn Research Center and Pratt & Whitney, a ceramic heat exchanger panel intended for use along the hot-flow-path walls of future reusable launch vehicles was designed, fabricated, and tested. These regeneratively cooled ceramic matrix composite (CMC) panels offer lighter weight, higher operating temperatures, and reduced coolant requirements in comparison to their more traditional metallic counterparts. A maintainable approach to the design was adopted which allowed the panel components to be assembled with high-temperature fasteners rather than by permanent bonding methods. With this approach, the CMC hot face sheet, the coolant containment system, and backside structure were all fabricated separately and could be replaced individually as the need occurred during use. This maintainable design leads to both ease of fabrication and reduced cost.

  1. Simulation of High-Speed Cavity Flows in a Scramjet Engine by the Space-Time CESE Method

    DTIC Science & Technology

    2005-09-01

    considered here share a common flow velocity u. Let the total specific energy def U 2 E = e+- (2.33) 2 11 Letd),, i = 1,2,..., N, be the net mass of...x and the y directions, respectively; p is the static pressure; e = E + ½ (u2 + v2) is the specific total energy with E as the specific internal...p.u. 2, and the total energy by p.uCOc2 . The subscript 0o denotes the free stream condition. The cavity depth d is used as the length scale, and the

  2. Hyper-X Research Vehicle (HXRV) Experimental Aerodynamics Test Program Overview

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.; Woods, William C.; Engelund, Walter C.

    2000-01-01

    This paper provides an overview of the experimental aerodynamics test program to ensure mission success for the autonomous flight of the Hyper-X Research Vehicle (HXRV). The HXRV is a 12-ft long, 2700 lb lifting body technology demonstrator designed to flight demonstrate for the first time a fully airframe integrated scramjet propulsion system. Three flights are currently planned, two at Mach 7 and one at Mach 10, beginning in the fall of 2000. The research vehicles will be boosted to the prescribed scramjet engine test point where they will separate from the booster, stabilize, and initiate engine test. Following 5+ seconds of powered flight and 15 seconds of cow-open tares, the cowl will close and the vehicle will fly a controlled deceleration trajectory which includes numerous control doublets for in-flight aerodynamic parameter identification. This paper reviews the preflight testing activities, wind tunnel models, test rationale, risk reduction activities, and sample results from wind tunnel tests supporting the flight trajectory of the HXRV from hypersonic engine test point through subsonic flight termination.

  3. HIFiRE Direct-Connect Rig (HDCR) Phase I Ground Test Results from the NASA Langley Arc-Heated Scramjet Test Facility

    NASA Technical Reports Server (NTRS)

    Hass, Neal E.; Cabell, Karen F.; Storch, Andrea M.

    2010-01-01

    The initial phase of hydrocarbon-fueled ground tests supporting Flight 2 of the Hypersonic International Flight Research Experiment (HIFiRE) Program has been conducted in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF). The HIFiRE Program, an Air Force-lead international cooperative program includes eight different flight test experiments designed to target specific challenges of hypersonic flight. The second of the eight planned flight experiments is a hydrocarbon-fueled scramjet flight test intended to demonstrate dual-mode to scramjet-mode operation and verify the scramjet performance prediction and design tools. A performance goal is the achievement of a combusted fuel equivalence ratio greater than 0.7 while in scramjet mode. The ground test rig, designated the HIFiRE Direct Connect Rig (HDCR), is a full-scale, heat sink, direct-connect ground test article that duplicates both the flowpath lines and the instrumentation layout of the isolator and combustor portion of the flight test hardware. The primary objectives of the HDCR Phase I tests are to verify the operability of the HIFiRE isolator/combustor across the Mach 6.0-8.0 flight regime and to establish a fuel distribution schedule to ensure a successful mode transition prior to the HiFIRE payload Critical Design Review. Although the phase I test plans include testing over the Mach 6 to 8 flight simulation range, only Mach 6 testing will be reported in this paper. Experimental results presented here include flowpath surface pressure, temperature, and heat flux distributions that demonstrate the operation of the flowpath over a small range of test conditions around the nominal Mach 6 simulation, as well as a range of fuel equivalence ratios and fuel injection distributions. Both ethylene and a mixture of ethylene and methane (planned for flight) were tested. Maximum back pressure and flameholding limits, as well as a baseline fuel schedule, that covers the Mach 5.84-6.5 test space have been

  4. Simultaneous Measurements of Temperature and Major Species Concentration in a Hydrocarbon-Fueled Dual Mode Scramjet Using WIDECARS

    NASA Astrophysics Data System (ADS)

    Gallo, Emanuela Carolina Angela

    Width increased dual-pump enhanced coherent anti-Stokes Raman spectroscopy (WIDECARS) measurements were conducted in a McKenna air-ethylene premixed burner, at nominal equivalence ratio range between 0.55 and 2.50 to provide quantitative measurements of six major combustion species (C2H 4, N2, O2, H2, CO, CO2) concentration and temperature simultaneously. The purpose of this test was to investigate the uncertainties in the experimental and spectral modeling methods in preparation for an subsequent scramjet C2H4/air combustion test at the University of Virginia-Aerospace Research Laboratory. A broadband Pyrromethene (PM) PM597 and PM650 dye laser mixture and optical cavity were studied and optimized to excite the Raman shift of all the target species. Two hundred single shot recorded spectra were processed, theoretically fitted and then compared to computational models, to verify where chemical equilibrium or adiabatic condition occurred, providing experimental flame location and formation, species concentrations, temperature, and heat losses inputs to computational kinetic models. The Stark effect, temperature, and concentration errors are discussed. Subsequently, WIDECARS measurements of a premixed air-ethylene flame were successfully acquired in a direct connect small-scale dual-mode scramjet combustor, at University of Virginia Supersonic Combustion Facility (UVaSCF). A nominal Mach 5 flight condition was simulated (stagnation pressure p0 = 300 kPa, temperature T0 = 1200 K, equivalence ratio range ER = 0.3 -- 0.4). The purpose of this test was to provide quantitative measurements of the six major combustion species concentration and temperature. Point-wise measurements were taken by mapping four two-dimensional orthogonal planes (before, within, and two planes after the cavity flame holder) with respect to the combustor freestream direction. Two hundred single shot recorded spectra were processed and theoretically fitted. Mean flow and standard deviation are

  5. Large-Eddy / Reynolds-Averaged Navier-Stokes Simulations of a Dual-Mode Scramjet Combustor

    NASA Technical Reports Server (NTRS)

    Fulton, Jesse A.; Edwards, Jack R.; Hassan, Hassan A.; Rockwell, Robert; Goyne, Christopher; McDaniel, James; Smith, Chad; Cutler, Andrew; Johansen, Craig; Danehy, Paul M.; Kouchi, Toshinori

    2012-01-01

    Numerical simulations of reacting and non-reacting flows within a scramjet combustor configuration experimentally mapped at the University of Virginia s Scramjet Combustion Facility (operating with Configuration A ) are described in this paper. Reynolds-Averaged Navier-Stokes (RANS) and hybrid Large Eddy Simulation / Reynolds-Averaged Navier-Stokes (LES / RANS) methods are utilized, with the intent of comparing essentially blind predictions with results from non-intrusive flow-field measurement methods including coherent anti-Stokes Raman spectroscopy (CARS), hydroxyl radical planar laser-induced fluorescence (OH-PLIF), stereoscopic particle image velocimetry (SPIV), wavelength modulation spectroscopy (WMS), and focusing Schlieren. NC State's REACTMB solver was used both for RANS and LES / RANS, along with a 9-species, 19- reaction H2-air kinetics mechanism by Jachimowski. Inviscid fluxes were evaluated using Edwards LDFSS flux-splitting scheme, and the Menter BSL turbulence model was utilized in both full-domain RANS simulations and as the unsteady RANS portion of the LES / RANS closure. Simulations were executed and compared with experiment at two equivalence ratios, PHI = 0.17 and PHI = 0.34. Results show that the PHI = 0.17 flame is hotter near the injector while the PHI = 0.34 flame is displaced further downstream in the combustor, though it is still anchored to the injector. Reactant mixing was predicted to be much better at the lower equivalence ratio. The LES / RANS model appears to predict lower overall heat release compared to RANS (at least for PHI = 0.17), and its capability to capture the direct effects of larger turbulent eddies leads to much better predictions of reactant mixing and combustion in the flame stabilization region downstream of the fuel injector. Numerical results from the LES/RANS model also show very good agreement with OH-PLIF and SPIV measurements. An un-damped long-wave oscillation of the pre-combustion shock train, which caused

  6. X-43C Plans and Status

    NASA Technical Reports Server (NTRS)

    Moses, Paul L.

    2003-01-01

    X-43C Project is a hypersonic flight demonstration being executed as a collaboration between the National Aeronautics and Space Administration (NASA) and the United States Air Force (USAF). X-43C will expand the hypersonic flight envelope for air breathing engines beyond the history making efforts of the Hyper-X Program (X-43A). X-43C will demonstrate sustained accelerating flight during three flight tests of expendable X-43C Demonstrator Vehicles (DVs). The approximately 16-foot long X-43C DV will be boosted to the starting test conditions, separate from the booster, and accelerate from Mach 5 to Mach 7 under its own power and autonomous control. The DVs are to be powered by a liquid hydrocarbon-fueled, fuel-cooled, dual-mode, airframe integrated scramjet engine system developed under the USAF HyTech Program. The Project is managed by NASA Langley Research Center as part of NASA s Next Generation Launch Technology Program. Flight tests will be conducted by NASA Dryden Flight Research Center over water off the coast of California in the Pacific Test Range. The NASA/USAF/industry project is a natural extension of the Hyper-X Program (X-43A), which will demonstrate short duration ( 10 seconds) gaseous hydrogen-fueled scramjet powered flight at Mach 7 and Mach 10 using a heavyweight, largely heat sink construction, experimental engine. The X-43C Project will demonstrate sustained accelerating flight from Mach 5 to Mach 7 ( 4 minutes) using a flight-weight, fuel-cooled, scramjet engine powered by much denser liquid hydrocarbon fuel. The X-43C DV design flows from integrating USAF HyTech developed engine technologies with a NASA Air Breathing Launch Vehicle accelerator-class configuration and Hyper-X heritage vehicle systems designs. This paper describes the X-43C Project and provides background for NASA s current hypersonic flight demonstration efforts.

  7. X-43C Flight Demonstrator Project Overview

    NASA Technical Reports Server (NTRS)

    Moses, Paul L.

    2003-01-01

    The X-43C Flight Demonstrator Project is a joint NASA-USAF hypersonic propulsion technology flight demonstration project that will expand the hypersonic flight envelope for air-breathing engines. The Project will demonstrate sustained accelerating flight through three flights of expendable X-43C Demonstrator Vehicles (DVs). The approximately 16-foot long X-43C DV will be boosted to the starting test conditions, separate from the booster, and accelerate from Mach 5 to Mach 7 under its own power and autonomous control. The DVs will be powered by a liquid hydrocarbon-fueled, fuel-cooled, dual-mode, airframe integrated scramjet engine system developed under the USAF HyTech Program. The Project is managed by NASA Langley Research Center as part of NASA's Next Generation Launch Technology Program. Flight tests will be conducted by NASA Dryden Flight Research Center off the coast of California over water in the Pacific Test Range. The NASA/USAF/industry project is a natural extension of the Hyper-X Program (X-43A), which will demonstrate short duration (approximately 10 seconds) gaseous hydrogen-fueled scramjet powered flight at Mach 7 and Mach 10 using a heavy-weight, largely heat sink construction, experimental engine. The X-43C Project will demonstrate sustained accelerating flight from Mach 5 to Mach 7 (approximately 4 minutes) using a flight-weight, fuel-cooled, scramjet engine powered by much denser liquid hydrocarbon fuel. The X-43C DV design flows from integrating USAF HyTech developed engine technologies with a NASA Air-Breathing Launch Vehicle accelerator-class configuration and Hyper-X heritage vehicle systems designs. This paper describes the X-43C Project and provides the background for NASA's current hypersonic flight demonstration efforts.

  8. Supersonic Flow Choking in Engine Wind Tunnels

    NASA Astrophysics Data System (ADS)

    Mitani, Tohru; Miyajima, Hiroshi; Tani, Koichiro; Kouchi, Toshinori; Sakuranaka, Noboru; Watanabe, Syuichi

    Breakdown of diffuser flow was often observed in our scramjet engine tests. This facility operation may damage the engine wind-tunnel and should be prevented. An one-dimensional analysis was applied to the diffuser flow to identify the causes of the flow breakdown. All the losses and gains by engine and friction loss in the diffuser were represented by point-sources of mass, momentum and energy. The thermal choking condition was calculated by uses of a chemical equilibrium code. The fuel rates causing the flow-choking successfully reproduced the limit fuel rates observed in our tests. Inlet-unstart of engine lost the ejector-pumping effect in the diffuser system to trigger the flow choking. The choking was also promoted by the drag of the gas sampling rakes. The choking in diffuser flow and the engine unstart may couple each other to cause hysteresis in the diffuser breakdown, which was also experienced in our tests. A rocket-based, combined-cycle (RBCC) engine will be tested under the Mach 4 condition. The engine easily causes the choking of diffuser because of the large propellant supply rates and the relatively-low specific impulse. Operation of the wind-tunnel was discussed to control the flow choking in the tests.

  9. NASA's Hypersonic Research Engine Project: A review

    NASA Technical Reports Server (NTRS)

    Andrews, Earl H.; Mackley, Ernest A.

    1994-01-01

    The goals of the NASA Hypersonic Research Engine (HRE) Project, which began in 1964, were to design, develop, and construct a high-performance hypersonic research ramjet/scramjet engine for flight tests of the developed concept over the speed range of Mach 4 to 8. The project was planned to be accomplished in three phases: project definition, research engine development, and flight test using the X-15A-2 research airplane, which was modified to carry hydrogen fuel for the research engine. The project goal of an engine flight test was eliminated when the X-15 program was canceled in 1968. Ground tests of full-scale engine models then became the focus of the project. Two axisymmetric full-scale engine models, having 18-inch-diameter cowls, were fabricated and tested: a structural model and combustion/propulsion model. A brief historical review of the project, with salient features, typical data results, and lessons learned, is presented. An extensive number of documents were generated during the HRE Project and are listed.

  10. Wind-tunnel blockage and actuation systems test of a two-dimensional scramjet inlet unstart model at Mach 6

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1994-01-01

    The present study examines the wind-tunnel blockage and actuation systems effectiveness in starting and forcibly unstarting a two-dimensional scramjet inlet in the NASA Langley 20-Inch Mach 6 Tunnel. The intent of the overall test program is to study (both experimentally and computationally) the dynamics of the inlet unstart; however, prior to the design and fabrication of an expensive, instrumented wind-tunnel model, it was deemed necessary first to examine potential wind-tunnel blockage issues related to model sizing and to examine the adequacy of the actuation systems in accomplishing the start and unstart. The model is equipped with both a moveable cowl and aft plug. Windows in the inlet sidewalls allow limited optical access to the internal shock structure; schlieren video was used to identify inlet start and unstart. A chronology of each actuation sequence is provided in tabular form along with still frames from the schlieren video. A pitot probe monitored the freestream conditions throughout the start/unstart process to determine if there was a blockage effect due to the model start or unstart. Because the purpose of this report is to make the phase I (blockage and actuation systems) data rapidly available to the community, the data is presented largely without analysis of the internal shock interactions or the unstart process. This series of tests indicated that the model was appropriately sized for this facility and identified operability limits required first to allow the inlet to start and second to force the unstart.

  11. Experimental investigations on active cooling thermal protection structure of hydrocarbon-fueled scramjet combustor in arc heated facility

    NASA Astrophysics Data System (ADS)

    Jianqiang, Tu; Jinlong, Peng; Xianning, Yang; Lianzhong, Chen

    2016-10-01

    The active cooling thermal protection technology is the efficient method to resolve the long-duration work and reusable problems of hydrocarbon-fueled scramjet combustor, where worst thermo-mechanical loads occur. The fuel is passed through coolant channels adjacent to the heated surfaces to absorb heat from the heating exchanger panels, prior to injection into the combustor. The heating exchanger both cooled down the wall temperature of the combustor wall and heats and cracks the hydrocarbon fuel inside the panel to permit an easier combustion and satisfying combustion efficiency. The subscale active cooling metallic panels, with dimensions of 100×100 mm and different coolant channel sizes, have been tested under typical combustion thermal environment produced by arc heated Turbulent Flow Duct (TFD). The heat exchange ability of different coolant channel sizes has been obtained. The big-scale active cooling metallic panel, with dimensions of 100 × 750 mm and the coolant channel sizes of better heating exchange performance, has been made and tested in the big-scale arc heated TFD facility. The test results show that the local superheated ablation is easy to happen for the cooling fuel assigned asymmetrically in the bigscale active cooling metallic panel, and the cooling fuel rate can reduce 8%˜10% after spraying the Thermal Barrier Coating (TBC) in the heating surface.

  12. Experimental investigation of generic three-dimensional sidewall-compression scramjet inlets at Mach 6 in tetrafluoromethane

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1993-01-01

    Three-dimensional sidewall-compression scramjet inlets with leading-edge sweeps of 30 deg and 70 deg were tested in the Langley Hypersonic CF4 Tunnel at Mach 6 and with a ratio of specific heats of 1.2. The parametric effects of leading-edge sweep, cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with 42 static pressure orifices that were distributed on the sidewalls, base plate, and cowl. Schlieren movies were made of each test for flow visualization of the effects of the internal flow spillage on the external flow field. To obtain an approximate characterization of the flow field, a modification to two-dimensional, inviscid, oblique shock theory was derived to accommodate the three-dimensional effects of leading-edge sweep. This theory qualitatively predicted the reflected shock structure (i.e., sidewall impingement locations) and the observed increase in spillage with increasing leading-edge sweep. The primary effect of moving the cowl forward was capturing the flow that would have otherwise spilled out ahead of the cowl. Increasing the contraction ratio increases the number of internal shock reflections and hence incrementally increases the sidewall pressure distribution. Significant Reynolds number effects were noted over a small range of Reynolds number.

  13. 1998 Calibration of the Mach 4.7 and Mach 6 Arc-Heated Scramjet Test Facility Nozzles

    NASA Technical Reports Server (NTRS)

    Witte, David W.; Irby, Richard G.; Auslender, Aaron H.; Rock, Kenneth E.

    2004-01-01

    A calibration of the Arc-Heated Scramjet Test Facility (AHSTF) Mach 4.7 and Mach 6 nozzles was performed in 1998. For each nozzle, three different typical facility operating test points were selected for calibration. Each survey consisted of measurements, at 340 separate locations across the 11 inch square nozzle exit plane, of pitot pressure, static pressure, and total temperature. Measurement density was higher (4/inch) in the boundary layer near the nozzle wall than in the core nozzle flow (1/inch). The results generated for each of these calibration surveys were contour plots at the nozzle exit plane of the measured and calculated flow properties which completely defined the thermodynamic state of the nozzle exit flow. An area integration of the mass flux at the nozzle exit for each survey was compared to the AHSTF mass flow meter results to provide an indication of the overall quality of the calibration performed. The percent difference between the integrated nozzle exit mass flow and the flow meter ranged from 0.0 to 1.3 percent for the six surveys. Finally, a comparison of this 1998 calibration was made with the 1986 calibration. Differences of less than 10 percent were found within the nozzle core flow while in the boundary layer differences on the order of 20 percent were quite common.

  14. Scramjet Development Tests Supporting the Mach 10 Flight of the X-43

    NASA Technical Reports Server (NTRS)

    Rogers, R. C.; Shih, A. T.; Hass, N. E.

    2005-01-01

    The Hyper-X Project s successful third flight of the X-43 at near Mach 10 in 2004 proved the potential for airbreathing propulsion at hypersonic speeds. The engine flowpath used in the X-43 research vehicle was developed and evaluated in a systematic series of ground tests in the NASA HyPulse Shock Tunnel at conditions duplicating Mach 10 flight using a full scale height, partial width engine model of the flight engine. Tests were conducted over a range of equivalence ratios from 0.8 to 1.6 using hydrogen and a mixture of two-percent silane in hydrogen fuels. Silane gas was used as an ignition aid during the short duration of the pulse facility tests. Variation of the engine inflow conditions, pressure, temperature, and Mach number, were parametrically varied during the test entries to broaden the database over the expected uncertainty in the flight conditions. A review of the ground test technique and comparisons of the ground test pressures along with selected flight data are presented.

  15. Generation After Next Propulsor Research: Robust Design for Embedded Engine Systems

    NASA Technical Reports Server (NTRS)

    Arend, David J.; Tillman, Gregory; O'Brien, Walter F.

    2012-01-01

    The National Aeronautics and Space Administration, United Technologies Research Center and Virginia Polytechnic and State University have contracted to pursue multi-disciplinary research into boundary layer ingesting (BLI) propulsors for generation after next environmentally responsible subsonic fixed wing aircraft. This Robust Design for Embedded Engine Systems project first conducted a high-level vehicle system study based on a large commercial transport class hybrid wing body aircraft, which determined that a 3 to 5 percent reduction in fuel burn could be achieved over a 7,500 nanometer mission. Both pylon-mounted baseline and BLI propulsion systems were based on a low-pressure-ratio fan (1.35) in an ultra-high-bypass ratio engine (16), consistent with the next generation of advanced commercial turbofans. An optimized, coupled BLI inlet and fan system was subsequently designed to achieve performance targets identified in the system study. The resulting system possesses an inlet with total pressure losses less than 0.5%, and a fan stage with an efficiency debit of less than 1.5 percent relative to the pylon-mounted, clean-inflow baseline. The subject research project has identified tools and methodologies necessary for the design of next-generation, highly-airframe-integrated propulsion systems. These tools will be validated in future large-scale testing of the BLI inlet / fan system in NASA's 8 foot x 6 foot transonic wind tunnel. In addition, fan unsteady response to screen-generated total pressure distortion is being characterized experimentally in a JT15D engine test rig. These data will document engine sensitivities to distortion magnitude and spatial distribution, providing early insight into key physical processes that will control BLI propulsor design.

  16. Enriched Storable Oxidizers for Rocket Engines

    NASA Technical Reports Server (NTRS)

    Sackheim, R. L.; Herdy, J. R., Jr.

    2010-01-01

    The name "enriched storable oxidizers" (ESOs) has been coined for a family of optimized mixtures of between two and four oxidizer fluids. For most applications, the constituents of these mixtures would be nitrogen tetroxide (N2O4), nitrous oxide (N2O), and nitrogen dioxide (NO2); in some applications, the mixtures might include inhibited red fuming nitric acid [IRFN (which consists of red fuming nitric acid to which some hydrogen fluoride is added to reduce its corrosive effect]. The optimum proportions of these constituents would be different for different applications. ESOs were originally proposed for use in spacecraft and launch-rocket propulsion systems: ESOs could be especially useful in advanced spacecraft propulsion systems that could operate in multiple modes. ESOs might also be useful in special terrestrial applications that could include ramjet and scramjet aircraft engines.

  17. Flight Control Laws for NASA's Hyper-X Research Vehicle

    NASA Technical Reports Server (NTRS)

    Davidson, J.; Lallman, F.; McMinn, J. D.; Martin, J.; Pahle, J.; Stephenson, M.; Selmon, J.; Bose, D.

    1999-01-01

    The goal of the Hyper-X program is to demonstrate and validate technology for design and performance predictions of hypersonic aircraft with an airframe-integrated supersonic-combustion ramjet propulsion system. Accomplishing this goal requires flight demonstration of a hydrogen-fueled scramjet powered hypersonic aircraft. A key enabling technology for this flight demonstration is flight controls. Closed-loop flight control is required to enable a successful stage separation, to achieve and maintain the design condition during the engine test, and to provide a controlled descent. Before the contract award, NASA developed preliminary flight control laws for the Hyper-X to evaluate the feasibility of the proposed scramjet test sequence and descent trajectory. After the contract award, a Boeing/NASA partnership worked to develop the current control laws. This paper presents a description of the Hyper-X Research Vehicle control law architectures with performance and robustness analyses. Assessments of simulated flight trajectories and stability margin analyses demonstrate that these control laws meet the flight test requirements.

  18. Rocketdyne Development of RBCC Engine for Low Cost Access to Space

    NASA Technical Reports Server (NTRS)

    Ortwerth, P.; Ratekin, G.; Goldman, A.; Emanuel, M.; Ketchum, A.; Horn, M.

    1997-01-01

    Rocketdyne is pursuing the conceptual design and development of a Rocket Based Combined Cycle (RBCC) engine for booster and SSTO, advanced reusable space transportation ARTT systems under contract with NASA Marshall Space Flight Center. The Rocketdyne concept is fixed geometry integrated Rocket, Ram Scramjet which is Hydrogen fueled and uses Hydrogen regenerative cooling. Vision vehicle integration studies have determined that scramjet operation to Mach 12 has high payoff for low cost reusable space transportation. Rocketdyne is internally developing versions of the concept for other applications in high speed aircraft and missiles with Hydrocarbon fuel systems. Subscale engine ground testing is underway for all modes of operation from takeoff to Mach 8. High altitude Rocket only mode tests will be completed as part of the ground test program to validate high expansion ratio performance. A unique feature of the ground test series is the inclusion of dynamic trajectory simulation with real time Mach number, altitude, engine throttling, and RBCC mode changes in a specially modified freejet test facility at GASL. Preliminary cold flow Air Augmented Rocket mode test results and Short Combustor tests have met program goals and have been used to integrate all modes of operation in a single combustor design with a fixed geometry inlet for design confirmation tests. A water cooled subscale engine is being fabricated and installed for test beginning the last quarter of 1997.

  19. Ultra-Efficient Engine Technology Project Continued to Contribute to Breakthrough Technologies

    NASA Technical Reports Server (NTRS)

    Shaw, Robert J.

    2003-01-01

    The International Civil Aviation Organization (ICAO), the U.S. Environmental Protection Agency (EPA), local environmental groups, and the public have become increasingly concerned over damage to local air quality from aircraft emissions and the impact of producing greenhouse gases. The NASA Glenn Research Center has been working to develop revolutionary technologies to minimize environmentally harmful engine emissions, such as nitrogen oxides, carbon dioxide, aerosols, and particulates. The two objectives of UEET are (1) to develop technologies to reduce nitrogen oxide (NOx) emissions by 70 percent below 1996 ICAO regulations and (2) to decrease carbon dioxide emissions (CO2) by dramatically increasing performance and efficiency. High temperature engine materials, ultra-low-NOx combustor designs, efficient, highly loaded turbomachinery, and propulsion-airframe integration analysis are technologies being developed at Glenn to meet these goals. Technology developed in the previous Advanced Subsonic Technology Program is being put into commercial production for large and regional aircraft to reduce NOx emissions 50 percent below 1996 ICAO regulations for landing and takeoff cycles. UEET will take the technology to the next quantum leap-reducing emissions to 70 percent below the ICAO regulations level. In addition, NASA-developed research will significantly reduce carbon monoxide, unburned hydrocarbons, and corresponding cruise NOx levels for the next generation of aircraft engines. Glenn's UEET research will be useful across the whole range of flight: subsonic, supersonic, and hypersonic. It will improve the subsonic transportation that the public depends on, contribute to supersonic commercial aircraft, improve military aircraft, and contribute to the design of a future hypersonic vehicle. These technologies are contributing to a better quality of life on Earth.

  20. Integrated Design and Engineering Analysis (IDEA) Environment - Propulsion Related Module Development and Vehicle Integration

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hilmi N.

    2013-01-01

    This report documents the work performed during the period from May 2011 - October 2012 on the Integrated Design and Engineering Analysis (IDEA) environment. IDEA is a collaborative environment based on an object-oriented, multidisciplinary, distributed framework using the Adaptive Modeling Language (AML). This report will focus on describing the work done in the areas of: (1) Integrating propulsion data (turbines, rockets, and scramjets) in the system, and using the data to perform trajectory analysis; (2) Developing a parametric packaging strategy for a hypersonic air breathing vehicles allowing for tank resizing when multiple fuels and/or oxidizer are part of the configuration; and (3) Vehicle scaling and closure strategies.

  1. High temperature dynamic engine seal technology development

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.; Dellacorte, Christopher; Machinchick, Michael; Mutharasan, Rajakkannu; Du, Guang-Wu; Ko, Frank; Sirocky, Paul J.; Miller, Jeffrey H.

    1992-01-01

    Combined cycle ramjet/scramjet engines being designed for advanced hypersonic vehicles, including the National Aerospace Plane (NASP), require innovative high temperature dynamic seals to seal the sliding interfaces of the articulated engine panels. New seals are required that will operate hot (1200 to 2000 F), seal pressures ranging from 0 to 100 psi, remain flexible to accommodate significant sidewall distortions, and resist abrasion over the engine's operational life. This report reviews the recent high temperature durability screening assessments of a new braided rope seal concept, braided of emerging high temperature materials, that shows promise of meeting many of the seal demands of hypersonic engines. The paper presents durability data for: (1) the fundamental seal building blocks, a range of candidate ceramic fiber tows; and for (2) braided rope seal subelements scrubbed under engine simulated sliding, temperature, and preload conditions. Seal material/architecture attributes and limitations are identified through the investigations performed. The paper summarizes the current seal technology development status and presents areas in which future work will be performed.

  2. Instrumentation requirements from the user's view. [For airbreathing hypersonic engines

    SciTech Connect

    Harsha, P.T.

    1988-01-01

    The use of combustor diagnostics is considered from the point of view of demonstration of performance of an airbreathing hypersonic engine. The basic need is seen to be that of providing the data necessary to verify performance predictions for the engine as installed in the airplane. This necessitates the use of a diagnostics capability that can provide the inputs required by the computational analyses that will be used to assess this performance. Because of the cost of ground test facilities, a premium is placed on measurement technique reliability and redundancy of instrumentation. A mix of nonintrusive optical techniques and probe-based measurements is seen to be the best approach using current diagnostics capability; one such instrument mix is outlined for a ramjet/scramjet test program. 11 references.

  3. International Symposium on Air Breathing Engines, 8th, Cincinnati, OH, June 14-19, 1987, Proceedings

    SciTech Connect

    Billig, F.S.

    1987-01-01

    The present conference on air-breathing aircraft engine technology considers topics in inlet design, radial-flow turbomachinery, fuel injection and combustion systems, axial flow compressor design and performance, ramjet configurations, turbine flow phenomena, engine control and service life, fluid flow-related problems, engine diagnostic methods, propfan design, combustor performance and pollutant chemistry, combustion dynamics, and engine system analysis. Attention is given to thrust-vectoring systems, supersonic missile air intakes, three-dimensional centrifugal compressors, airblast atomizers, secondary flows in axial flow compressors, axial compressor blade tip clearance flows, hydrogen scramjets with sidewall injection, the performance of a variable-geometry turbine, advanced tip clearance control systems, rotary jet mixing, fan blade aeroelastic behavior, flow dynamics in combustion processes, and the technology of low cost turbomachinery.

  4. Test Medium Effects on Ignition, Combustion and Flameholding Processes in Scramjet Combustors

    NASA Technical Reports Server (NTRS)

    Chinitz, W.; Pellett, G. L.

    2000-01-01

    The total temperatures (enthalpies) required to ground-test air-breathing (aero-propulsion) engines at high Mach number flight conditions can be achieved in a number of ways. Among these are: 1. Heat exchangers, including pre-heated ceramic beds. 2. direct electrical heating, e.g., arc discharge and resistance heaters. 3. Compression heating. 4. Shock heating, and 5. In-stream combustion, with oxygen replenishment to match air content. Each method has distinct advantages, disadvantages and limitations. All have a common characteristic of being designed for intermittent flow, due to the extreme energy required for continuous operation at simulated Mach numbers above about 3. All also distort the composition of atmospheric air to some degree, due to the high temperatures that occur in the plenum section prior to expansion of the flow to simulated flight conditions. In the case of in-stream combustion, the resulting test medium is commonly referred to as "vitiated air", being composed of oxygen, nitrogen and some fraction of combustion products.

  5. Wind-tunnel tests on a 3-dimensional fixed-geometry scramjet inlet at M = 2.30 to 4.60

    NASA Technical Reports Server (NTRS)

    Mueller, J. N.; Trexler, C. A.; Souders, S. W.

    1977-01-01

    Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.

  6. Recent Activities in Research of the Combined Cycle Engine at JAXA

    NASA Astrophysics Data System (ADS)

    Tani, Kouichiro; Tomioka, Sadatake; Kato, Kanenori; Ueda, Syuichi; Takegoshi, Masao

    Recent activities of the researches on the rocket based combined cycle engine in Japan Aerospace Exploration Agency are summarized. Aiming to realize the flight test in 10 years, JAXA has been making sub-scale model experiments as well as a series of component tests. In 08 fiscal year, sub-scale tests were carried out in Mach 6 flight condition and the stable ramjet combustion was confirmed following the successful ramjet mode establishment in Mach 4 condition in previous year. Some improvements of flow modeling inside the combustor and the ejector analysis were also achieved. With the scramjet mode analysis due in ’09 fiscal year, the designing method of the combined cycle engine will be improved and the next test engine will be launched.

  7. Testing of the NASA Hypersonics Project Combined Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LlMX)

    NASA Technical Reports Server (NTRS)

    Saunders, J. D.; Stueber, T. J.; Thomas, S. R.; Suder, K. L.; Weir, L. J.; Sanders, B. W.

    2012-01-01

    Status on an effort to develop Turbine Based Combined Cycle (TBCC) propulsion is described. This propulsion technology can enable reliable and reusable space launch systems. TBCC propulsion offers improved performance and safety over rocket propulsion. The potential to realize aircraft-like operations and reduced maintenance are additional benefits. Among most the critical TBCC enabling technologies are: 1) mode transition from turbine to scramjet propulsion, 2) high Mach turbine engines and 3) TBCC integration. To address these TBCC challenges, the effort is centered on a propulsion mode transition experiment and includes analytical research. The test program, the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LIMX), was conceived to integrate TBCC propulsion with proposed hypersonic vehicles. The goals address: (1) dual inlet operability and performance, (2) mode-transition sequences enabling a switch between turbine and scramjet flow paths, and (3) turbine engine transients during transition. Four test phases are planned from which a database can be used to both validate design and analysis codes and characterize operability and integration issues for TBCC propulsion. In this paper we discuss the research objectives, features of the CCE hardware and test plans, and status of the parametric inlet characterization testing which began in 2011. This effort is sponsored by the NASA Fundamental Aeronautics Hypersonics project

  8. The X-43A (Hyper-X) Flies Into the Record Books

    NASA Technical Reports Server (NTRS)

    Grindle, Laurie; Bahm, Catherine

    2006-01-01

    The goal of the Hyper-X research program, conducted jointly by the NASA Dryden Flight Research Center and the NASA Langley Research Center, was to demonstrate and validate the technology, experimental techniques, and computation methods and tools for design and performance predictions of a hypersonic aircraft with an airframe-integrated, scramjet propulsion system. Three X-43A airframe-integrated, scramjet research vehicles were designed and fabricated to achieve that goal by flight test: two test flights at Mach 7 and one test flight at Mach 10. The first flight, conducted on June 2, 2001, experienced a launch vehicle failure and resulted in a 9-month mishap investigation. A two-year return-to-flight effort ensued and concluded when the second Mach 7 flight was successful on March 27, 2004. Just eight months later, on November 16, the X-43A successfully completed the third and final flight. These two flights were the first flight demonstrations, at Mach 7 and Mach 10 respectively, of an airframe-integrated, scramjet-powered, hypersonic vehicle.

  9. Ongoing Analysis of Rocket Based Combined Cycle Engines by the Applied Fluid Dynamics Analysis Group at Marshall Space Flight Center

    NASA Technical Reports Server (NTRS)

    Ruf, Joseph; Holt, James B.; Canabal, Francisco

    1999-01-01

    This paper presents the status of analyses on three Rocket Based Combined Cycle configurations underway in the Applied Fluid Dynamics Analysis Group (TD64). TD64 is performing computational fluid dynamics analysis on a Penn State RBCC test rig, the proposed Draco axisymmetric RBCC engine and the Trailblazer engine. The intent of the analysis on the Penn State test rig is to benchmark the Finite Difference Navier Stokes code for ejector mode fluid dynamics. The Draco engine analysis is a trade study to determine the ejector mode performance as a function of three engine design variables. The Trailblazer analysis is to evaluate the nozzle performance in scramjet mode. Results to date of each analysis are presented.

  10. Advanced fabrication techniques for hydrogen-cooled engine structures

    NASA Technical Reports Server (NTRS)

    Buchmann, O. A.; Arefian, V. V.; Warren, H. A.; Vuigner, A. A.; Pohlman, M. J.

    1985-01-01

    Described is a program for development of coolant passage geometries, material systems, and joining processes that will produce long-life hydrogen-cooled structures for scramjet applications. Tests were performed to establish basic material properties, and samples constructed and evaluated to substantiate fabrication processes and inspection techniques. Results of the study show that the basic goal of increasing the life of hydrogen-cooled structures two orders of magnitude relative to that of the Hypersonic Research Engine can be reached with available means. Estimated life is 19000 cycles for the channels and 16000 cycles for pin-fin coolant passage configurations using Nickel 201. Additional research is required to establish the fatigue characteristics of dissimilar-metal coolant passages (Nickel 201/Inconel 718) and to investigate the embrittling effects of the hydrogen coolant.

  11. Rocket-Based Combined Cycle Engine Concept Development

    NASA Technical Reports Server (NTRS)

    Ratekin, G.; Goldman, Allen; Ortwerth, P.; Weisberg, S.; McArthur, J. Craig (Technical Monitor)

    2001-01-01

    The development of rocket-based combined cycle (RBCC) propulsion systems is part of a 12 year effort under both company funding and contract work. The concept is a fixed geometry integrated rocket, ramjet, scramjet, which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals, seal purge gas, and closeout side attachments. Engine A5 is the current configuration for NASA Marshall Space Flight Center (MSFC) for the ART program. Engine A5 models the complete flight engine flowpath of inlet, isolator, airbreathing combustor, and nozzle. High-performance rocket thrusters are integrated into the engine enabling both low speed air-augmented rocket (AAR) and high speed pure rocket operation. Engine A5 was tested in GASL's new Flight Acceleration Simulation Test (FAST) facility in all four operating modes, AAR, RAM, SCRAM, and Rocket. Additionally, transition from AAR to RAM and RAM to SCRAM was also demonstrated. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. SCRAM and rocket mode performance was above predictions. For the first time, testing also demonstrated transition between operating modes.

  12. Ongoing Analyses of Rocket Based Combined Cycle Engines by the Applied Fluid Dynamics Analysis Group at Marshall Space Flight Center

    NASA Technical Reports Server (NTRS)

    Ruf, Joseph H.; Holt, James B.; Canabal, Francisco

    2001-01-01

    This paper presents the status of analyses on three Rocket Based Combined Cycle (RBCC) configurations underway in the Applied Fluid Dynamics Analysis Group (TD64). TD64 is performing computational fluid dynamics (CFD) analysis on a Penn State RBCC test rig, the proposed Draco axisymmetric RBCC engine and the Trailblazer engine. The intent of the analysis on the Penn State test rig is to benchmark the Finite Difference Navier Stokes (FDNS) code for ejector mode fluid dynamics. The Draco analysis was a trade study to determine the ejector mode performance as a function of three engine design variables. The Trailblazer analysis is to evaluate the nozzle performance in scramjet mode. Results to date of each analysis are presented.

  13. Engineering Encounters: Engineering Adaptations

    ERIC Educational Resources Information Center

    Gatling, Anne; Vaughn, Meredith Houle

    2015-01-01

    Engineering is not a subject that has historically been taught in elementary schools, but with the emphasis on engineering in the "Next Generation Science Standards," curricula are being developed to explicitly teach engineering content and design. However, many of the scientific investigations already conducted with students have…

  14. Boundary Layer Transition on X-43A

    NASA Technical Reports Server (NTRS)

    Berry, Scott; Daryabeigi, Kamran; Wurster, Kathryn; Bittner, Robert

    2008-01-01

    The successful Mach 7 and 10 flights of the first fully integrated scramjet propulsion systems by the Hyper-X (X-43A) program have provided the means with which to verify the original design methodologies and assumptions. As part of Hyper-X s propulsion-airframe integration, the forebody was designed to include a spanwise array of vortex generators to promote boundary layer transition ahead of the engine. Turbulence at the inlet is thought to provide the most reliable engine design and allows direct scaling of flight results to groundbased data. Pre-flight estimations of boundary layer transition, for both Mach 7 and 10 flight conditions, suggested that forebody boundary layer trips were required to ensure fully turbulent conditions upstream of the inlet. This paper presents the results of an analysis of the thermocouple measurements used to infer the dynamics of the transition process during the trajectories for both flights, on both the lower surface (to assess trip performance) and the upper surface (to assess natural transition). The approach used in the analysis of the thermocouple data is outlined, along with a discussion of the calculated local flow properties that correspond to the transition events as identified in the flight data. The present analysis has confirmed that the boundary layer trips performed as expected for both flights, providing turbulent flow ahead of the inlet during critical portions of the trajectory, while the upper surface was laminar as predicted by the pre-flight analysis.

  15. Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Airbreathing Hypersonic Engines

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.; Bityurin, Valentine A.; Lineberry, John T.

    1999-01-01

    Established analyses of conventional ramjet/scramjet performance characteristics indicate that a considerable decrease in efficiency can be expected at off-design flight conditions. This can be explained, in large part, by the deterioration of intake mass flow and limited inlet compression at low flight speeds and by the onset of thrust degradation effects associated with increased burner entry temperature at high flight speeds. In combination, these effects tend to impose lower and upper Mach number limits for practical flight. It has been noted, however, that Magnetohydrodynamic (MHD) energy management techniques represent a possible means for extending the flight Mach number envelope of conventional engines. By transferring enthalpy between different stages of the engine cycle, it appears that the onset of thrust degradation may be delayed to higher flight speeds. Obviously, the introduction of additional process inefficiencies is inevitable with this approach, but it is believed that these losses are more than compensated through optimization of the combustion process. The fundamental idea is to use MHD energy conversion processes to extract and bypass a portion of the intake kinetic energy around the burner. We refer to this general class of propulsion system as an MHD-bypass engine. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass airbreathing hypersonic engines using ideal gasdynamics and fundamental thermodynamic principles.

  16. Schlieren photographs and internal pressure distributions for three-dimensional sidewall-compression scramjet inlets at a Mach number of 6 in CF4

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.

    1993-01-01

    Three-dimensional sidewall-compression scramjet inlets with leading-edge sweeps of 30 deg and 70 deg were tested in the Langley Hypersonic CF4 Tunnel at a Mach number of 6 and a free-stream ratio of specific heats of 1.2. The parametric effects of leading-edge sweep, cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with static pressure orifices distributed on the sidewalls, baseplate, and cowl. Schlieren movies were made of selected tunnel runs for flow visualization of the entrance plane and cowl region. Although these movies could not show the internal flow, the effect of the internal flow on the external flow was evident by way of spillage. The purpose is to provide a preliminary data release for the investigation. The models, facility, and testing methods are described, and the test matrix and a tabulation of tunnel runs are provided. Line plots highlighting the stated parametric effects and a representative set of schlieren photographs are presented without analysis.

  17. Performance of a hydrogen burner to simulate air entering scramjet combustors. [simulation of total temperature, total pressure, and volume fraction of oxygen of air at flight conditions

    NASA Technical Reports Server (NTRS)

    Russin, W. R.

    1974-01-01

    Tests were conducted to determine the performance of a hydrogen burner used to produce a test gas that simulates air entering a scramjet combustor at various flight conditions. The test gas simulates air in that it duplicates the total temperature, total pressure, and the volume fraction of oxygen of air at flight conditions. The main objective of the tests was to determine the performance of the burner as a function of the effective exhaust port area. The conclusions were: (1) pressure oscillations of the chugging type were reduced in amplitude to plus or minus 2 percent of the mean pressure level by proper sizing of hydrogen, oxygen, and air injector flow areas; (2) combustion efficiency remained essentially constant as the exhaust port area was increased by a factor of 3.4; (3) the mean total temperature determined from integrating the exit radial gas property profiles was within plus or minus 5 percent of the theoretical bulk total temperature; (4) the measured exit total temperature profile had a local peak temperature more than 30 percent greater than the theoretical bulk total temperature; and (5) measured heat transfer to the burner liner was 75 percent of that predicted by theory based on a flat radial temperature profile.

  18. Effect of variation of length-to-depth ratio and Mach number on the performance of a typical double cavity scramjet combustor

    NASA Astrophysics Data System (ADS)

    Mahto, Navin Kumar; Choubey, Gautam; Suneetha, Lakka; Pandey, K. M.

    2016-11-01

    The two equation standard k-ɛ turbulence model and the two-dimensional compressible Reynolds-Averaged Navier-Stokes (RANS) equations have been used to computationally simulate the double cavity scramjet combustor. Here all the simulations are performed by using ANSYS 14-FLUENT code. At the same time, the validation of the present numerical simulation for double cavity has been performed by comparing its result with the available experimental data which is in accordance with the literature. The results are in good agreement with the schlieren image and the pressure distribution curve obtained experimentally. However, the pressure distribution curve obtained numerically is under-predicted in 5 locations by numerical calculation. Further, investigations on the variations of the effects of the length-to-depth ratio of cavity and Mach number on the combustion characteristics has been carried out. The present results show that there is an optimal length-to-depth ratio for the cavity for which the performance of combustor significantly improves and also efficient combustion takes place within the combustor region. Also, the shifting of the location of incident oblique shock took place in the downstream of the H2 inlet when the Mach number value increases. But after achieving a critical Mach number range of 2-2.5, the further increase in Mach number results in lower combustion efficiency which may deteriorate the performance of combustor.

  19. An experimental parametric study of geometric, Reynolds number, and ratio of specific heats effects in three-dimensional sidewall compression scramjet inlets at Mach 6

    NASA Technical Reports Server (NTRS)

    Holland, Scott D.; Murphy, Kelly J.

    1993-01-01

    The effects of the decrease in the ratio of specific heats have been investigated in generic 3D sidewall compression scramjet inlets with leading-edge sweep angles of 30 and 70 degrees. The effects of a decrease in ratio of specific heats were seen by comparing data from two facilities in two test gases: in the Langley Mach 6 CF4 Tunnel in tetrafluoromethane and in the Langley 15-Inch Mach 6 Air Tunnel in perfect gas air. In addition to the simulated real gas effects, the parametric effects of cowl position, contraction ratio, leading-edge sweep, and Reynolds number were investigated in the 15-Inch Mach 6 Air Tunnel. The models were instrumented with a total of 45 static pressure orifices distributed on the sidewalls and baseplate. Surface streamline patterns were examined via oil flow, and schlieren videos were made of the external flow field. The results of these tests have significant implications to ground based testing of inlets in facilities which do not operate at flight enthalpies.

  20. High temperature performance evaluation of a hypersonic engine ceramic wafer seal

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.

    1991-01-01

    Leakage rates of an innovative hypersonic engine seal were measured using a specially developed static high temperature seal test fixture at NASA Lewis Research Center. The three foot long structural panel-edge seal is designed to minimize leakage of high temperature, high pressure gases past the movable panels of advanced ramjet/scramjet engines. The seal is made of a stack of precision machined ceramic wafer pieces that are inserted into a closely conforming seal channel in the movable engine panel. The wafer seal accommodates the significant distortions in the adjacent engine walls through relative sliding between adjacent wafers. Seal leakage rates are presented for engine simulated air temperatures up to 1350 F and for engine pressures up to 100 psi. Leakage rates are also presented for the seal, sealing both a flat wall condition, and an engine simulated distorted wall condition in which the distortion was 0.15 in. in only an 18 in. span. Seal leakage rates were low, meeting an industry-established tentative leakage limit for all combinations of temperature, pressure, and wall conditions considered. Comparisons are made between the measured leakage rates and leakage rates predicted using a seal leakage model developed from externally-pressurized gas film bearing theory.

  1. High temperature performance evaluation of a hypersonic engine ceramic wafer seal

    SciTech Connect

    Steinetz, B.M.

    1991-04-01

    Leakage rates of an innovative hypersonic engine seal were measured using a specially developed static high temperature seal test fixture at NASA Lewis Research Center. The three foot long structural panel-edge seal is designed to minimize leakage of high temperature, high pressure gases past the movable panels of advanced ramjet/scramjet engines. The seal is made of a stack of precision machined ceramic wafer pieces that are inserted into a closely conforming seal channel in the movable engine panel. The wafer seal accommodates the significant distortions in the adjacent engine walls through relative sliding between adjacent wafers. Seal leakage rates are presented for engine simulated air temperatures up to 1350F and for engine pressures up to 100 psi. Leakage rates are also presented for the seal, sealing both a flat wall condition, and an engine simulated distorted wall condition in which the distortion was 0.15 in. in only an 18 in. span. Seal leakage rates were low, meeting an industry-established tentative leakage limit for all combinations of temperature, pressure, and wall conditions considered. Comparisons are made between the measured leakage rates and leakage rates predicted using a seal leakage model developed from externally-pressurized gas film bearing theory.

  2. Options for flight testing rocket-based combined-cycle (RBCC) engines

    NASA Technical Reports Server (NTRS)

    Olds, John

    1996-01-01

    While NASA's current next-generation launch vehicle research has largely focused on advanced all-rocket single-stage-to-orbit vehicles (i.e. the X-33 and it's RLV operational follow-on), some attention is being given to advanced propulsion concepts suitable for 'next-generation-and-a-half' vehicles. Rocket-based combined-cycle (RBCC) engines combining rocket and airbreathing elements are one candidate concept. Preliminary RBCC engine development was undertaken by the United States in the 1960's. However, additional ground and flight research is required to bring the engine to technological maturity. This paper presents two options for flight testing early versions of the RBCC ejector scramjet engine. The first option mounts a single RBCC engine module to the X-34 air-launched technology testbed for test flights up to about Mach 6.4. The second option links RBCC engine testing to the simultaneous development of a small-payload (220 lb.) two-stage-to-orbit operational vehicle in the Bantam payload class. This launcher/testbed concept has been dubbed the W vehicle. The W vehicle can also serve as an early ejector ramjet RBCC launcher (albeit at a lower payload). To complement current RBCC ground testing efforts, both flight test engines will use earth-storable propellants for their RBCC rocket primaries and hydrocarbon fuel for their airbreathing modes. Performance and vehicle sizing results are presented for both options.

  3. Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

    NASA Astrophysics Data System (ADS)

    Forde, Scott; Bulman, Mel; Neill, Todd

    2006-07-01

    Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

  4. The Hyper-X Flight Systems Validation Program

    NASA Technical Reports Server (NTRS)

    Redifer, Matthew; Lin, Yohan; Bessent, Courtney Amos; Barklow, Carole

    2007-01-01

    For the Hyper-X/X-43A program, the development of a comprehensive validation test plan played an integral part in the success of the mission. The goal was to demonstrate hypersonic propulsion technologies by flight testing an airframe-integrated scramjet engine. Preparation for flight involved both verification and validation testing. By definition, verification is the process of assuring that the product meets design requirements; whereas validation is the process of assuring that the design meets mission requirements for the intended environment. This report presents an overview of the program with emphasis on the validation efforts. It includes topics such as hardware-in-the-loop, failure modes and effects, aircraft-in-the-loop, plugs-out, power characterization, antenna pattern, integration, combined systems, captive carry, and flight testing. Where applicable, test results are also discussed. The report provides a brief description of the flight systems onboard the X-43A research vehicle and an introduction to the ground support equipment required to execute the validation plan. The intent is to provide validation concepts that are applicable to current, follow-on, and next generation vehicles that share the hybrid spacecraft and aircraft characteristics of the Hyper-X vehicle.

  5. Pitot survey of exhaust flow field of a 2-D scramjet nozzle at Mach 6 with air or freon and argon used for exhaust simulation

    NASA Technical Reports Server (NTRS)

    Monta, William J.

    1992-01-01

    A pitot-rake survey of the simulated exhaust of a half-span scramjet nozzle model was conducted in the Langley 20-Inch Mach 6 Tunnel to provide an additional data set for computational fluid dynamics (CFD) code comparisons. A wind-tunnel model was tested with a 26-tube pitot rake that could be manually positioned along the mid-semispan plane of the model. The model configuration had an external expansion surface of 20 degrees and an internal cowl expansion of 12 degrees; tests were also performed with a flow fence. Tests were conducted at a free-stream Reynolds number of approximately 6.5 x 10(exp 6) per foot and a model angle of attack of -0.75 degrees. The two exhaust gas mediums that were tested were air and a Freon 12-argon mixture. Each medium was tested at two jet total pressures at approximately 28 and 14 psia. This document presents the flow-field survey results in graphical as well as tabular form, and several observations concerning the results are discussed. The surveys reveal the major expected flow-field characteristics for each test configuration. For a 50-percent freon 12 and 50-percent argon mixture by volume (Fr-Ar), the exhaust jet pressures were slightly higher than those for air. The addition of a flow fence slightly raised the pitot pressure for the Fr-Ar mixture, but it produced little change for air. For the Fr-Ar exhaust, the plume was larger and the region between the shock wave and plume was smaller.

  6. A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy; Alexander, Reginald

    1999-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

  7. Advanced fabrication techniques for hydrogen-cooled engine structures. Final report, October 1975-June 1982

    SciTech Connect

    Buchmann, O.A.; Arefian, V.V.; Warren, H.A.; Vuigner, A.A.; Pohlman, M.J.

    1985-11-01

    Described is a program for development of coolant passage geometries, material systems, and joining processes that will produce long-life hydrogen-cooled structures for scramjet applications. Tests were performed to establish basic material properties, and samples constructed and evaluated to substantiate fabrication processes and inspection techniques. Results of the study show that the basic goal of increasing the life of hydrogen-cooled structures two orders of magnitude relative to that of the Hypersonic Research Engine can be reached with available means. Estimated life is 19000 cycles for the channels and 16000 cycles for pin-fin coolant passage configurations using Nickel 201. Additional research is required to establish the fatigue characteristics of dissimilar-metal coolant passages (Nickel 201/Inconel 718) and to investigate the embrittling effects of the hydrogen coolant.

  8. Orbit-on-demand vehicle propelled by air-turborocket/ramjet engines

    NASA Technical Reports Server (NTRS)

    Hartung, L.; Karkow, J.; Ordway, W.; Pickett, D.; Muras, A.

    1986-01-01

    A preliminary design study has been completed for a fully reusable, single-stage-to-orbit transatmospheric vehicle. The specified mission capability was to lift a 20,000 lb payload to low earth orbit. A ground accelerator-assisted horizontal take-off was chosen to increase operational flexibility. The multi-mode propulsion system included the use of air-turborocket, ramjet, scramjet and rocket engines. Weight and performance estimates were obtained for the vehicle. A computer package was developed to perform aerothermodynamic analyses of the propulsion modes throughout the flight environment from take-off to low earth orbit. Results are presented for a semi-optimized trajectory. The analysis indicates that a vehicle of this type has great potential for providing low cost, flexible access to space.

  9. Orbit-on-demand vehicle propelled by air-turborocket/ramjet engines

    NASA Astrophysics Data System (ADS)

    Hartung, L.; Karkow, J.; Ordway, W.; Pickett, D.; Muras, A.

    1986-06-01

    A preliminary design study has been completed for a fully reusable, single-stage-to-orbit transatmospheric vehicle. The specified mission capability was to lift a 20,000 lb payload to low earth orbit. A ground accelerator-assisted horizontal take-off was chosen to increase operational flexibility. The multi-mode propulsion system included the use of air-turborocket, ramjet, scramjet and rocket engines. Weight and performance estimates were obtained for the vehicle. A computer package was developed to perform aerothermodynamic analyses of the propulsion modes throughout the flight environment from take-off to low earth orbit. Results are presented for a semi-optimized trajectory. The analysis indicates that a vehicle of this type has great potential for providing low cost, flexible access to space.

  10. Experimental investigation of inlet-combustor isolators for a dual-mode scramjet at a Mach number of 4

    NASA Technical Reports Server (NTRS)

    Emami, Saied; Trexler, Carl A.; Auslender, Aaron H.; Weidner, John P.

    1995-01-01

    This report details experimentally derived operational characteristics of numerous two-dimensional planar inlet-combustor isolator configurations at a Mach number of 4. Variations in geometry included (1) inlet cowl length; (2) inlet cowl rotation angle; (3) isolator length; and (4) utilization of a rearward-facing isolator step. To obtain inlet-isolator maximum pressure-rise data relevant to ramjet-engine combustion operation, configurations were mechanically back pressured. Results demonstrated that the combined inlet-isolator maximum back-pressure capability increases as a function of isolator length and contraction ratio, and that the initiation of unstart is nearly independent of inlet cowl length, inlet cowl contraction ratio, and mass capture. Additionally, data are presented quantifying the initiation of inlet unstarts and the corresponding unstart pressure levels.

  11. Layered Systems Engineering Engines

    NASA Technical Reports Server (NTRS)

    Breidenthal, Julian C.; Overman, Marvin J.

    2009-01-01

    A notation is described for depicting the relationships between multiple, contemporaneous systems engineering efforts undertaken within a multi-layer system-of-systems hierarchy. We combined the concepts of remoteness of activity from the end customer, depiction of activity on a timeline, and data flow to create a new kind of diagram which we call a "Layered Vee Diagram." This notation is an advance over previous notations because it is able to be simultaneously precise about activity, level of granularity, product exchanges, and timing; these advances provide systems engineering managers a significantly improved ability to express and understand the relationships between many systems engineering efforts. Using the new notation, we obtain a key insight into the relationship between project duration and the strategy selected for chaining the systems engineering effort between layers, as well as insights into the costs, opportunities, and risks associated with alternate chaining strategies.

  12. Unsteady pressure loads in a generic high speed engine model

    NASA Technical Reports Server (NTRS)

    Parrott, Tony L.; Jones, Michael G.; Thurlow, Ernie M.

    1992-01-01

    Unsteady pressure loads were measured along the top interior wall of a generic high-speed engine (GHSE) model undergoing performance tests in the combustion-Heated Scramjet Test Facility at the Langley Research Center. Flow to the model inlet was simulated at 72000 ft and a flight Mach number of 4. The inlet Mach number was 3.5 with a total temperature and pressure of 1640 R and 92 psia. The unsteady pressure loads were measured with 5 piezoresistive gages, recessed into the wall 4 to 12 gage diameters to reduce incident heat flux to the diaphragms, and distributed from the inlet to the combustor. Contributors to the unsteady pressure loads included boundary layer turbulence, combustion noise, and transients generated by unstart loads. Typical turbulent boundary layer rms pressures in the inlet ranged from 133 dB in the inlet to 181 dB in the combustor over the frequency range from 0 to 5 kHz. Downstream of the inlet exist, combustion noise was shown to dominate boundary layer turbulence noise at increased heat release rates. Noise levels in the isolator section increased by 15 dB when the fuel-air ratio was increased from 0.37 to 0.57 of the stoichiometric ratio. Transient pressure disturbances associated with engine unstarts were measured in the inlet and have an upstream propagation speed of about 7 ft/sec and pressure jumps of at least 3 psia.

  13. Development of braided rope seals for hypersonic engine applications: Flow modeling

    NASA Technical Reports Server (NTRS)

    Mutharasan, Rajakkannu; Steinetz, Bruce M.; Tao, Xiaoming; Du, Guang-Wu; Ko, Frank

    1992-01-01

    A new type of engine seal is being developed to meet the needs of advanced hypersonic engines. A seal braided of emerging high temperature ceramic fibers comprised of a sheath-core construction was selected for study based on its low leakage rates. Flexible, low-leakage, high temperature seals are required to seal the movable engine panels of advanced ramjet-scramjet engines either preventing potentially dangerous leakage into backside engine cavities or limiting the purge coolant flow rates through the seals. To predict the leakage through these flexible, porous seal structures new analytical flow models are required. Two such models based on the Kozeny-Carman equations are developed herein and are compared to experimental leakage measurements for simulated pressure and seal gap conditions. The models developed allow prediction of the gas leakage rate as a function of fiber diameter, fiber packing density, gas properties, and pressure drop across the seal. The first model treats the seal as a homogeneous fiber bed. The second model divides the seal into two homogeneous fiber beds identified as the core and the sheath of the seal. Flow resistances of each of the main seal elements are combined to determine the total flow resistance. Comparisons between measured leakage rates and model predictions for seal structures covering a wide range of braid architectures show good agreement. Within the experimental range, the second model provides a prediction within 6 to 13 percent of the flow for many of the cases examined. Areas where future model refinements are required are identified.

  14. A new cascade-less engine operated from subsonic to hypersonic conditions: designed by computational fluid dynamics of compressible turbulence with chemical reactions

    NASA Astrophysics Data System (ADS)

    Naitoh, Ken; Nakamura, Kazushi; Emoto, Takehiro

    2010-12-01

    By using our computational fluid dynamic models, a new type of single engine capable of operating over a wide range of Mach numbers from subsonic to hypersonic regimes is proposed for airplanes, whereas traditional piston engines, turbojet engines, and scram engines work only under a narrower range of operating conditions. The new engine has no compressors or turbines such as those used in conventional turbojet engines. An important point is its system of super multijets that collide to compress gas for the transonic regime. Computational fluid dynamics is applied to clarify the potential of this engine. The peak pressure at the combustion center is over 2.5 MPa, while that just before ignition is over 1.0 MPa. The maximum power of this engine will be sufficient for actual use. Under the conditions of higher Mach numbers, the main intake passage located in front of the super multijet nozzles, takes in air more. That results in a ram or scramjet engine for supersonic and hypersonic conditions.

  15. Engineering Practice and Engineering Ethics.

    ERIC Educational Resources Information Center

    Lynch, William T.; Kline, Ronald

    2000-01-01

    Offers ways of applying science and technology studies to the teaching of engineering ethics. Suggests modifications of both detailed case studies on engineering disasters and hypothetical, ethical dilemmas employed in engineering ethics classes. (Author/CCM)

  16. Stirling engines

    SciTech Connect

    Reader, G.T.; Hooper

    1983-01-01

    The Stirling engine was invented by a Scottish clergyman in 1816, but fell into disuse with the coming of the diesel engine. Advances in materials science and the energy crisis have made a hot air engine economically attractive. Explanations are full and understandable. Includes coverage of the underlying thermodynamics and an interesting historical section. Topics include: Introduction to Stirling engine technology, Theoretical concepts--practical realities, Analysis, simulation and design, Practical aspects, Some alternative energy sources, Present research and development, Stirling engine literature.

  17. Neural Engineering

    NASA Astrophysics Data System (ADS)

    He, Bin

    About the Series: Bioelectric Engineering presents state-of-the-art discussions on modern biomedical engineering with respect to applications of electrical engineering and information technology in biomedicine. This focus affirms Springer's commitment to publishing important reviews of the broadest interest to biomedical engineers, bioengineers, and their colleagues in affiliated disciplines. Recent volumes have covered modeling and imaging of bioelectric activity, neural engineering, biosignal processing, bionanotechnology, among other topics.

  18. Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

    NASA Technical Reports Server (NTRS)

    Felder, James L.; Kim, Hyun Dae; Brown, Gerald V.

    2009-01-01

    possibilities. The Boeing N2 hybrid-wing-body (HWB) is used as a baseline aircraft for this study. The two pylon mounted conventional turbofans are replaced by two wing-tip mounted turboshaft engines, each driving a superconducting generator. Both generators feed a common electrical bus which distributes power to an array of superconducting motor-driven fans in a continuous nacelle centered along the trailing edge of the upper surface of the wing-body. A key finding was that traditional inlet performance methodology has to be modified when most of the air entering the inlet is boundary layer air. A very thorough and detailed propulsion/airframe integration (PAI) analysis is required at the very beginning of the design process since embedded engine inlet performance must be based on conditions at the inlet lip rather than freestream conditions. Examination of a range of fan pressure ratios yielded a minimum Thrust-specific-fuel-consumption (TSFC) at the aerodynamic design point of the vehicle (31,000 ft /Mach 0.8) between 1.3 and 1.35 FPR. We deduced that this was due to the higher pressure losses prior to the fan inlet as well as higher losses in the 2-D inlets and nozzles. This FPR is likely to be higher than the FPR that yields a minimum TSFC in a pylon mounted engine. 1

  19. Optical engineering

    SciTech Connect

    Saito, T T

    1998-01-01

    The Optical Engineering thrust area at Lawrence Livermore National Laboratory (LLNL) was created in the summer of 1996 with the following main objectives: (1) to foster and stimulate leading edge optical engineering research and efforts key to carrying out LLNL's mission and enabling major new programs; (2) to bring together LLNL's broad spectrum of high level optical engineering expertise to support its programs. Optical engineering has become a pervasive and key discipline, with applications across an extremely wide range of technologies, spanning the initial conception through the engineering refinements to enhance revolutionary application. It overlaps other technologies and LLNL engineering thrust areas.

  20. Hypersonic research engine project. Phase 2: Preliminary report on the performance of the HRE/AIM at Mach 6

    NASA Technical Reports Server (NTRS)

    Sun, Y. H.; Sainio, W. C.

    1975-01-01

    Test results of the Aerothermodynamic Integration Model are presented. A program was initiated to develop a hydrogen-fueled research-oriented scramjet for operation between Mach 3 and 8. The primary objectives were to investigate the internal aerothermodynamic characteristics of the engine, to provide realistic design parameters for future hypersonic engine development as well as to evaluate the ground test facility and testing techniques. The engine was tested at the NASA hypersonic tunnel facility with synthetic air at Mach 5, 6, and 7. The hydrogen fuel was heated up to 1500 R prior to injection to simulate a regeneratively cooled system. The engine and component performance at Mach 6 is reported. Inlet performance compared very well both with theory and with subscale model tests. Combustor efficiencies up to 95 percent were attained at an equivalence ratio of unity. Nozzle performance was lower than expected. The overall engine performance was computed using two different methods. The performance was also compared with test data from other sources.

  1. Engineering Motion

    ERIC Educational Resources Information Center

    Tuttle, Nicole; Stanley, Wendy; Bieniek, Tracy

    2016-01-01

    For many teachers, engineering can be intimidating; teachers receive little training in engineering, particularly those teaching early elementary students. In addition, the necessity of differentiating for students with special needs can make engineering more challenging to teach. This article describes a professional development program…

  2. Technologies for Propelled Hypersonic Flight Volume 2 - Subgroup 2: Scram Propulsion

    DTIC Science & Technology

    2006-01-01

    builds upon earlier co-operation in defence research and technology as set-up under the Advisory Group for Aerospace Research and Development (AGARD...solid base for the future. The content of this publication has been reproduced directly from material supplied by RTO or the authors. Published...Airframe Integration 1-17 1.5 Materials and Structures 1-18 1.6 Flight Testing 1-20 1.6.1 Hyper-X 1-20 1.6.2 CIAM – “Kholod” Scramjet

  3. Development and Implementation of a Scramjet Cycle Analysis Code with a Finite-Rate-Chemistry Combustion Model for Use on a Personal Computer

    DTIC Science & Technology

    1993-12-01

    structural and material engineers, since many minor species can be extremely corrosive ( Askeland , 1984: 670). 2.2.2. - Equilibrium and Frozen Flow...York: McGraw-Hill, 1989. Angelucci, E. la of Military Airaft. New York: Crescent Books, 1990. Askeland , D.R. The Science and Engineering of Materials

  4. Analytical and experimental investigations of the oblique detonation wave engine concept

    NASA Technical Reports Server (NTRS)

    Menees, Gene P.; Adelman, Henry G.; Cambier, Jean-Luc

    1990-01-01

    Wave combustors, which include the oblique detonation wave engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using Computational Fluid Dynamics (CFD) codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being performed in an arc heated hypersonic wind tunnel. Several fuel injection design were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

  5. Analytical and experimental investigations of the oblique detonation wave engine concept

    NASA Technical Reports Server (NTRS)

    Menees, Gene P.; Adelman, Henry G.; Cambier, Jean-Luc

    1991-01-01

    Wave combustors, which include the Oblique Detonation Wave Engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which will require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using CFD codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being carried out in an arc heated hypersonic wind tunnel. Several fuel injection designs were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

  6. Engine Lubricant

    NASA Technical Reports Server (NTRS)

    1993-01-01

    PS 212, a plasma-sprayed coating developed by NASA, is used to coat valves in a new rotorcam engine. The coating eliminates the need for a liquid lubricant in the rotorcam, which has no crankshaft, flywheel, distributor or water pump. Developed by Murray United Development Corporation, it is a rotary engine only 10 inches long with four cylinders radiating outward from a central axle. Company officials say the engine will be lighter, more compact and cheaper to manufacture than current engines and will feature cleaner exhaust emissions. A licensing arrangement with a manufacturer is under negotiation. Primary applications are for automobiles, but the engine may also be used in light aircraft.

  7. Shockwave Engine: Wave Disk Engine

    SciTech Connect

    2010-01-14

    Broad Funding Opportunity Announcement Project: MSU is developing a new engine for use in hybrid automobiles that could significantly reduce fuel waste and improve engine efficiency. In a traditional internal combustion engine, air and fuel are ignited, creating high-temperature and high-pressure gases which expand rapidly. This expansion of gases forces the engine’s pistons to pump and powers the car. MSU’s engine has no pistons. It uses the combustion of air and fuel to build up pressure within the engine, generating a shockwave that blasts hot gas exhaust into the blades of the engine’s rotors causing them to turn, which generates electricity. MSU’s redesigned engine would be the size of a cooking pot and contain fewer moving parts—reducing the weight of the engine by 30%. It would also enable a vehicle that could use 60% of its fuel for propulsion.

  8. Research and Development of Ram/Scramjets and Turboramjets in Russia (La Recherche et le Developpement des Statoreacteurs, des Statoreacteurs a Combustion Supersonique et des Turbostatoreacteurs en Russie)

    DTIC Science & Technology

    1993-12-01

    complex includes of stationary supersonic wind tunnels , hypersonic facilities of short-time action, impulse tun- nels, ballistic facilities and...of 130-ton class |I0|. Two liquid boosters were installed under missile’s delta wing . The ramjet engine, designed by M.M. Bondarynk. was integrated...parameters (thrust to total weight ratio, wing loading, engine static thrust per unit inlet area, inlet to wing area ratio) Hie modern mathematical

  9. Invisible Engineers

    NASA Astrophysics Data System (ADS)

    Ohashi, Hideo

    Questionnaire to ask “mention three names of scientists you know” and “three names of engineers you know” was conducted and the answers from 140 adults were analyzed. The results indicated that the image of scientists is represented by Nobel laureates and that of engineers by great inventors like Thomas Edison and industry founders like Soichiro Honda. In order to reveal the image of engineers among young generation, questionnaire was conducted for pupils in middle and high schools. Answers from 1,230 pupils were analyzed and 226 names mentioned as engineers were classified. White votes reached 60%. Engineers who are neither big inventors nor company founders collected less than 1% of named votes. Engineers are astonishingly invisible from young generation. Countermeasures are proposed.

  10. Aerojet Engine

    NASA Technical Reports Server (NTRS)

    1999-01-01

    NASA is developing technology for air-breathing rocket engines that could help make space transportation safe, reliable and affordable for ordinary people. Powered by engines that breathe oxygen from the air, the spacecraft would be completely reusable, take off and land at airport runways, and be ready to fly again within days. The engines would get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engines rely totally on oxygen in the atmosphere to burn the hydrogen fuel. Once the vehicle's speed increases to about 10 times the speed of sound, the engine converts to a conventional rocket-power mode to propel the vehicle into orbit. This Quick Time movie features an aerojet engine which is the main propulsion system of the X-series future launch vehicles.

  11. Information engineering

    SciTech Connect

    Hunt, D.N.

    1997-02-01

    The Information Engineering thrust area develops information technology to support the programmatic needs of Lawrence Livermore National Laboratory`s Engineering Directorate. Progress in five programmatic areas are described in separate reports contained herein. These are entitled Three-dimensional Object Creation, Manipulation, and Transport, Zephyr:A Secure Internet-Based Process to Streamline Engineering Procurements, Subcarrier Multiplexing: Optical Network Demonstrations, Parallel Optical Interconnect Technology Demonstration, and Intelligent Automation Architecture.

  12. A Collaborative Analysis Tool for Integrated Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy; Alexander, Reginald; Landrum, Brian

    2000-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. Then there is a transition to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scramjet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance. To adequately determine the performance of the engine/vehicle, the Hypersonic Flight Inlet Model (HYFIM) module was designed to interface with the RBCC

  13. Engineering ethics beyond engineers' ethics.

    PubMed

    Basart, Josep M; Serra, Montse

    2013-03-01

    Engineering ethics is usually focused on engineers' ethics, engineers acting as individuals. Certainly, these professionals play a central role in the matter, but engineers are not a singularity inside engineering; they exist and operate as a part of a complex network of mutual relationships between many other people, organizations and groups. When engineering ethics and engineers' ethics are taken as one and the same thing the paradigm of the ethical engineer which prevails is that of the heroic engineer, a certain model of the ideal engineer: someone both quite individualistic and strong enough to deal with all the moral challenges that could arise. We argue that this is not the best approach, at least today in our interrelated world. We have achieved a high degree of independence from nature by means of technology. In exchange for this autonomy we have become increasingly tied up with very complex systems to which we constantly delegate new tasks and powers. Concerns about safety keep growing everywhere due to the fact that now we have a sensitive awareness of the huge amount of power we are both consuming and deploying, thus, new forms of dialogue and consensus have to be incorporated at different levels, in different forums and at different times. Within these democratic channels of participation not just the needs and interests, but also the responsibilities and mutual commitments of all parties should be taken into account.

  14. Holistic Engineering

    ERIC Educational Resources Information Center

    Grasso, Domenico; Martinelli, David

    2007-01-01

    In this article, the authors discuss how to prepare high-quality engineers who are better equipped to serve in the changing global marketplace, and suggest educators in pursuing the holistic concept of the "unity of knowledge" that will yield a definition of engineering more fitting for the times ahead. The unity of knowledge is fundamentally…

  15. Engineering Administration.

    ERIC Educational Resources Information Center

    Naval Personnel Program Support Activity, Washington, DC.

    This book is intended to acquaint naval engineering officers with their duties in the engineering department. Standard shipboard organizations are analyzed in connection with personnel assignments, division operations, and watch systems. Detailed descriptions are included for the administration of directives, ship's bills, damage control, training…

  16. Corrosion Engineering.

    ERIC Educational Resources Information Center

    White, Charles V.

    A description is provided for a Corrosion and Corrosion Control course offered in the Continuing Engineering Education Program at the General Motors Institute (GMI). GMI is a small cooperative engineering school of approximately 2,000 students who alternate between six-week periods of academic study and six weeks of related work experience in…

  17. Electrochemical Engineering.

    ERIC Educational Resources Information Center

    Alkire, Richard C.

    1983-01-01

    Discusses engineering ramifications of electrochemistry, focusing on current/potential distribution, evaluation of trade-offs between influences of different phenomena, use of dimensionless numbers to assist in scale-over to new operating conditions, and economics. Also provides examples of electrochemical engineering education content related to…

  18. Systems Engineering

    NASA Technical Reports Server (NTRS)

    Pellerano, Fernando

    2015-01-01

    This short course provides information on what systems engineering is and how the systems engineer guides requirements, interfaces with the discipline leads, and resolves technical issues. There are many system-wide issues that either impact or are impacted by the thermal subsystem. This course will introduce these issues and illustrate them with real life examples.

  19. Genetic Engineering

    ERIC Educational Resources Information Center

    Phillips, John

    1973-01-01

    Presents a review of genetic engineering, in which the genotypes of plants and animals (including human genotypes) may be manipulated for the benefit of the human species. Discusses associated problems and solutions and provides an extensive bibliography of literature relating to genetic engineering. (JR)

  20. The X-43A hypersonic research aircraft and its modified Pegasus booster rocket mounted to NASA's NB

    NASA Technical Reports Server (NTRS)

    2001-01-01

    The first of three X-43A hypersonic research aircraft and its modified Pegasus booster rocket recently underwent combined systems testing while mounted to NASA's NB-52B carrier aircraft at the Dryden Flight Research Center, Edwards, California. The combined systems test was one of the last major milestones in the Hyper-X research program before the first X-43A flight. One of the major goals of the Hyper-X program is flight validation of airframe-integrated, air-breathing propulsion system, which so far have only been tested in ground facilities, such as wind tunnels. The X-43A flights will be the first actual flight tests of an aircraft powered by a revolutionary supersonic-combustion ramjet ('scramjet') engine capable of operating at hypersonic speeds above Mach 5 (five times the speed of sound). The X-43A design uses the underbody of the aircraft to form critical elements of the engine. The forebody shape helps compress the intake airflow, while the aft section acts as a nozzle to direct thrust. The 12-foot, unpiloted research vehicle was developed and built by MicroCraft Inc., Tullahoma, Tenn., under NASA contract. The booster, built by Orbital Sciences Corp., Dulles, Va., will accelerate the X-43A after the X-43A/booster 'stack' is air-launched from NASA's venerable NB-52 mothership. The X-43A will separate from the rocket at a predetermined altitude and speed and fly a pre-programmed trajectory, conducting aerodynamic and propulsion experiments until it descends into the Pacific Ocean. Three research flights are planned, two at Mach 7 and one at Mach 10.

  1. Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

    NASA Technical Reports Server (NTRS)

    Morgan, Morris H.; Gilinsky, Mikhail; Patel, Kaushal; Coston, Calvin; Blankson, Isaiah M.

    2003-01-01

    The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. Results obtained are based on analytical methods, numerical simulations and experimental tests at the NASA LaRC and Hampton University computer complexes and experimental facilities. The main objective of this research is injection, mixing and combustion enhancement in propulsion systems. The sub-projects in the reporting period are: (A) Aero-performance and acoustics of Telescope-shaped designs. The work included a pylon set application for SCRAMJET. (B) An analysis of sharp-edged nozzle exit designs for effective fuel injection into the flow stream in air-breathing engines: triangular-round and diamond-round nozzles. (C) Measurement technique improvements for the HU Low Speed Wind Tunnel (HU LSWT) including an automatic data acquisition system and a two component (drag-lift) balance system. In addition, a course in the field of aerodynamics was developed for the teaching and training of HU students.

  2. Analysis of the Magneto-Hydrodynamic (MHD) Energy Bypass Engine for High-Speed Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Riggins, David W.

    2002-01-01

    shock systems, finite-rate chemistry, wall cooling with thermally balanced engine (fuel heat sink), fuel injection and mixing, friction, etc. are shown and discussed for both the MHD engine and the conventional scramjet. The MHD bypass engine has significantly lower performance in all categories across the Mach number range (8 to 12.2). The lower performance is attributed to the combined effects of 1) additional irreversibility and cooling requirements associated with the MHD components and 2) the total pressure decrease associated with the inverse cycle itself.

  3. Engineering Performance

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Under Goddard Space Flight Center and Jet Propulsion Laboratory SBIR (Small Business Innovative Research) contracts, AST Engineering Services, Inc. developed a system engineering computer software tool to model how certain applications will affect a proposed system's performance. Quantitative System Engineering (QASE) evaluates system timing, capacity and availability. The system is used to predict performance of proposed real-time, aerospace systems, embedded systems, and/or scientific systems, as well as in support of NASA's EDOS (Earth Observing System (EOS) Data and Operations System) initiative.

  4. Green Engineering

    EPA Pesticide Factsheets

    Green Engineering is the design, commercialization and use of processes and products that are feasible and economical while reducing the generation of pollution at the source and minimizing the risk to human health and the environment.

  5. Harmonic engine

    DOEpatents

    Bennett, Charles L.

    2009-10-20

    A high efficiency harmonic engine based on a resonantly reciprocating piston expander that extracts work from heat and pressurizes working fluid in a reciprocating piston compressor. The engine preferably includes harmonic oscillator valves capable of oscillating at a resonant frequency for controlling the flow of working fluid into and out of the expander, and also preferably includes a shunt line connecting an expansion chamber of the expander to a buffer chamber of the expander for minimizing pressure variations in the fluidic circuit of the engine. The engine is especially designed to operate with very high temperature input to the expander and very low temperature input to the compressor, to produce very high thermal conversion efficiency.

  6. Electrochemical Engineering

    ERIC Educational Resources Information Center

    Alkire, Richard

    1976-01-01

    Discusses an electrochemical engineering course that combines transport phenomena and basic physical chemistry. Lecture notes and homework problems are used instead of a textbook; an outline of lecture topics is presented. (MLH)

  7. Engineering Geology

    ERIC Educational Resources Information Center

    Lee, Fitzhugh T.

    1974-01-01

    Briefly reviews the increasing application of geologic principles, techniques and data to engineering practices in the areas of land use and zoning controls, resource management energy programs and other fields. (BR)

  8. Flight-Ready TDLAS Combustion Sensor for the HIFiRE 2 Hypersonic Research Program

    DTIC Science & Technology

    2009-09-01

    environment near a scramjet combustor . The sensor should also be suitable as a diagnostic in realistic scramjet test situations, including in flight...4b column averaging 10 4c column-averaged absorption spectrum, 1395.45 - 1395.75 nm 11 5 jet engine combustor in-situ sensor 12 6 augmented jet...vibrational environment near a scramjet combustor . The sensor should also be suitable as a diagnostic in realistic scramjet test situations, including

  9. 2007 Research and Engineering Annual Report

    NASA Technical Reports Server (NTRS)

    Stoliker, Patrick; Bowers, Albion; Cruciani, Everlyn

    2008-01-01

    Selected research and technology activities at NASA Dryden Flight Research Center are summarized. These following activities exemplify the Center's varied and productive research efforts: Developing a Requirements Development Guide for an Automatic Ground Collision Avoidance System; Digital Terrain Data Compression and Rendering for Automatic Ground Collision Avoidance Systems; Nonlinear Flutter/Limit Cycle Oscillations Prediction Tool; Nonlinear System Identification Using Orthonormal Bases: Application to Aeroelastic/Aeroservoelastic Systems; Critical Aerodynamic Flow Feature Indicators: Towards Application with the Aerostructures Test Wing; Multidisciplinary Design, Analysis, and Optimization Tool Development Using a Genetic Algorithm; Structural Model Tuning Capability in an Object-Oriented Multidisciplinary Design, Analysis, and Optimization Tool; Extension of Ko Straight-Beam Displacement Theory to the Deformed Shape Predictions of Curved Structures; F-15B with Phoenix Missile and Pylon Assembly--Drag Force Estimation; Mass Property Testing of Phoenix Missile Hypersonic Testbed Hardware; ARMD Hypersonics Project Materials and Structures: Testing of Scramjet Thermal Protection System Concepts; High-Temperature Modal Survey of the Ruddervator Subcomponent Test Article; ARMD Hypersonics Project Materials and Structures: C/SiC Ruddervator Subcomponent Test and Analysis Task; Ground Vibration Testing and Model Correlation of the Phoenix Missile Hypersonic Testbed; Phoenix Missile Hypersonic Testbed: Performance Design and Analysis; Crew Exploration Vehicle Launch Abort System-Pad Abort-1 (PA-1) Flight Test; Testing the Orion (Crew Exploration Vehicle) Launch Abort System-Ascent Abort-1 (AA-1) Flight Test; SOFIA Flight-Test Flutter Prediction Methodology; SOFIA Closed-Door Aerodynamic Analyses; SOFIA Handling Qualities Evaluation for Closed-Door Operations; C-17 Support of IRAC Engine Model Development; Current Capabilities and Future Upgrade Plans of the C-17 Data

  10. Software engineering

    NASA Technical Reports Server (NTRS)

    Fridge, Ernest M., III; Hiott, Jim; Golej, Jim; Plumb, Allan

    1993-01-01

    Today's software systems generally use obsolete technology, are not integrated properly with other software systems, and are difficult and costly to maintain. The discipline of reverse engineering is becoming prominent as organizations try to move their systems up to more modern and maintainable technology in a cost effective manner. The Johnson Space Center (JSC) created a significant set of tools to develop and maintain FORTRAN and C code during development of the space shuttle. This tool set forms the basis for an integrated environment to reengineer existing code into modern software engineering structures which are then easier and less costly to maintain and which allow a fairly straightforward translation into other target languages. The environment will support these structures and practices even in areas where the language definition and compilers do not enforce good software engineering. The knowledge and data captured using the reverse engineering tools is passed to standard forward engineering tools to redesign or perform major upgrades to software systems in a much more cost effective manner than using older technologies. The latest release of the environment was in Feb. 1992.

  11. Rotary engine

    SciTech Connect

    Meyman, U.

    1987-02-03

    A rotary engine is described comprising: two covers spaced from one another; rotors located between the covers and rotating and planetating in different phases; the rotors interengaging to form working chambers therebetween; means to supply fluid to the working chambers and means to exhaust fluid from the working chambers during the operating cycle of the engine; gearing for synchronizing rotation and planetation of the rotors and each including first and second gears arranged so that one of the gears is connected with the rotors while the other of the gears is connected with an immovable part of the engine and the gears engage with one another; carriers interconnecting the rotors and planetating in the same phase with the planetation of the rotors for synchronizing the rotation and planetation of the rotors; shafts arranged to support the carriers during their planetations; and elements for connecting the covers with one another.

  12. Photoreceptor engineering

    PubMed Central

    Ziegler, Thea; Möglich, Andreas

    2015-01-01

    Sensory photoreceptors not only control diverse adaptive responses in Nature, but as light-regulated actuators they also provide the foundation for optogenetics, the non-invasive and spatiotemporally precise manipulation of cellular events by light. Novel photoreceptors have been engineered that establish control by light over manifold biological processes previously inaccessible to optogenetic intervention. Recently, photoreceptor engineering has witnessed a rapid development, and light-regulated actuators for the perturbation of a plethora of cellular events are now available. Here, we review fundamental principles of photoreceptors and light-regulated allostery. Photoreceptors dichotomize into associating receptors that alter their oligomeric state as part of light-regulated allostery and non-associating receptors that do not. A survey of engineered photoreceptors pinpoints light-regulated association reactions and order-disorder transitions as particularly powerful and versatile design principles. Photochromic photoreceptors that are bidirectionally toggled by two light colors augur enhanced spatiotemporal resolution and use as photoactivatable fluorophores. By identifying desirable traits in engineered photoreceptors, we provide pointers for the design of future, light-regulated actuators. PMID:26137467

  13. Harmonic engine

    DOEpatents

    Bennett, Charles L.; Sewall, Noel; Boroa, Carl

    2014-08-19

    An engine based on a reciprocating piston engine that extracts work from pressurized working fluid. The engine includes a harmonic oscillator inlet valve capable of oscillating at a resonant frequency for controlling the flow of working fluid into of the engine. In particular, the inlet valve includes an inlet valve head and a spring arranged together as a harmonic oscillator so that the inlet valve head is moveable from an unbiased equilibrium position to a biased closed position occluding an inlet. Upon releasing the inlet valve the inlet valve head undergoes a single oscillation past the equilibrium positio to a maximum open position and returns to a biased return position close to the closed position to choke the flow and produce a pressure drop across the inlet valve causing the inlet valve to close. Protrusions carried either by the inlet valve head or piston head are used to bump open the inlet valve from the closed position and initiate the single oscillation of the inlet valve head, and protrusions carried either by the outlet valve head or piston head are used to close the outlet valve ahead of the bump opening of the inlet valve.

  14. Engineering Geology

    ERIC Educational Resources Information Center

    Hatheway, Allen W.

    1978-01-01

    Engineering geology remains a potpourri of applied classical geology, and 1977 witnessed an upswing in demand for these services. Traditional foundation-related work was slight, but construction related to national needs increased briskly. Major cities turned to concerns of transit waste-water treatment and solid-waste disposal. (Author/MA)

  15. Adaptive Engineering

    ERIC Educational Resources Information Center

    VanderSteen, Jonathan

    2011-01-01

    Engineers today cannot meet their professional obligation to the welfare of society if they do not have a broad, multidisciplinary vision, and yet a multidisciplinary vision is becoming enormously difficult to obtain. A new curriculum must emerge that can integrate a focused, discipline-based scientific approach with an integrated approach. To do…

  16. Engineering seismology

    USGS Publications Warehouse

    N.N, Ambraseys

    1991-01-01

    Twenty years have elasped since the first issue of Earthquakes & Volcanoes. Apart from the remarkable increases in the number of scientists actively enagaged in earth sciences, what are the outstanding achievements during the past 20 years in the field of engineering seismology, which is my own speciality?

  17. Concurrent engineering

    NASA Technical Reports Server (NTRS)

    Chamis, C. C.; Leger, L.; Hunter, D.; Jones, C.; Sprague, R.; Berke, L.; Newell, J.; Singhal, S.

    1991-01-01

    The following subject areas are covered: issues (liquid rocket propulsion - current development approach, current certification process, and costs of engineering changes); state of the art (DICE information management system, key government participants, project development strategy, quality management, and numerical propulsion system simulation); needs identified; and proposed program.

  18. Engineering Students' Perceptions of Engineering Specialties

    ERIC Educational Resources Information Center

    Shivy, V.A.; Sullivan, T.N.

    2005-01-01

    The field of engineering is defined by a number of specialty areas, thus most engineering students must decide upon an educational specialty track within the engineering major. Data on familiarity with, and perceptions of similarity among 11 engineering specialties were collected from 129 undergraduate engineering students from a public urban…

  19. Software engineering as an engineering discipline

    NASA Technical Reports Server (NTRS)

    Berard, Edward V.

    1988-01-01

    The following topics are discussed in the context of software engineering: early use of the term; the 1968 NATO conference; Barry Boehm's definition; four requirements fo software engineering; and additional criteria for software engineering. Additionally, the four major requirements for software engineering--computer science, mathematics, engineering disciplines, and excellent communication skills--are discussed. The presentation is given in vugraph form.

  20. Enhancing Engineering Education through Engineering Management

    ERIC Educational Resources Information Center

    Pence, Kenneth R.; Rowe, Christopher J.

    2012-01-01

    Engineering Management courses are added to a traditional engineering curriculum to enhance the value of an undergraduate's engineering degree. A four-year engineering degree often leaves graduates lacking in business and management acumen. Engineering management education covers topics enhancing the value of new graduates by teaching management…

  1. Quiet engine program flight engine design study

    NASA Technical Reports Server (NTRS)

    Klapproth, J. F.; Neitzel, R. E.; Seeley, C. T.

    1974-01-01

    The results are presented of a preliminary flight engine design study based on the Quiet Engine Program high-bypass, low-noise turbofan engines. Engine configurations, weight, noise characteristics, and performance over a range of flight conditions typical of a subsonic transport aircraft were considered. High and low tip speed engines in various acoustically treated nacelle configurations were included.

  2. Engineering Review Information System

    NASA Technical Reports Server (NTRS)

    Grems, III, Edward G. (Inventor); Henze, James E. (Inventor); Bixby, Jonathan A. (Inventor); Roberts, Mark (Inventor); Mann, Thomas (Inventor)

    2015-01-01

    A disciplinal engineering review computer information system and method by defining a database of disciplinal engineering review process entities for an enterprise engineering program, opening a computer supported engineering item based upon the defined disciplinal engineering review process entities, managing a review of the opened engineering item according to the defined disciplinal engineering review process entities, and closing the opened engineering item according to the opened engineering item review.

  3. Combustion Characteristics of Liquid Normal Alkane Fuels in a Model Combustor of Supersonic Combustion Ramjet Engine

    NASA Astrophysics Data System (ADS)

    今村, 宰; 石川, 雄太; 鈴木, 俊介; 福本, 皓士郎; 西田, 俊介; 氏家, 康成; 津江, 光洋

    Effect of kinds of one-component n-alkane liquid fuels on combustion characteristics was investigated experimentally using a model combustor of scramjet engine. The inlet condition of a model combustor is 2.0 of Mach number, up to 2400K of total temperature, and 0.38MPa of total pressure. Five kinds of n-alkane are tested, of which carbon numbers are 7, 8, 10, 13, and 16. They are more chemically active and less volatile with an increase of alkane carbon number. Fuels are injected to the combustor in the upstream of cavity with barbotage nitrogen gas and self-ignition performance was investigated. The result shows that self-ignition occurs with less equivalence ratio when alkane carbon number is smaller. This indicates that physical characteristic of fuel, namely volatile of fuel, is dominant for self-ignition behavior. Effect on flame-holding performance is also examined with adding pilot hydrogen and combustion is kept after cutting off pilot hydrogen with the least equivalence ratio where alkane carbon number is from 8 to 10. These points are discussed qualitatively from the conflict effect of chemical and physical properties on alkane carbon number.

  4. Re-engineering Engineering Education

    ERIC Educational Resources Information Center

    Gordon, Bernard M.; Silevitch, Michael B.

    2009-01-01

    In 2005, leaders gathered by the National Association of Manufacturers declared yet another "STEM" emergency. In the face of global competition, they argued, the number of bachelor's degrees awarded annually to U.S. students in science, math and engineering must double by 2015. In fact, the need for STEM talent is even more critical…

  5. Heat engines

    NASA Astrophysics Data System (ADS)

    Rekos, N. F., Jr.; Parsons, E. L., Jr.

    1989-09-01

    For the past decade, the Department of Energy (DOE) has sponsored projects to develop diesel and gas turbine engines capable of operating on low-cost, coal-based fuels. Much of the current work addresses the use of coal-water fuel (CWF) in diesel and turbines, although there is some work with dry coal feed and other coal fuels. Both the diesel and gas turbine portions of the program include proof-of-concept and support projects. Specific highlights of the program include: engine tests and economic analyses have shown that CWF can replace 70 percent of the diesel oil used in the duty cycle of a typical main-line locomotive; A. D. Little and Cooper-Bessemer completed a system and economic study of coal-fueled diesel engines for modular power and industrial cogeneration markets. The coal-fueled diesel was found to be competitive at fuel oil prices of $5.50 per million British thermal units (MBtu); Over 200 hours of testing have been completed using CWF in full-scale, single-cylinder diesel engines. Combustion efficiencies have exceeded 99 percent; Both CWF and dry coal fuel forms can be burned in short residence time in-line combustors and in off-base combustors with a combustion efficiency of over 99 percent; Rich/lean combustion systems employed by the three major DOE contractors have demonstrated low NO(sub x) emissions levels; Contractors have also achieved promising results for controlling sulfur oxide (SO(sub x)) emissions using calcium-based sorbents; Slagging combustors have achieved between 65 and 95 percent slag capture, which will limit particulate loading on pre-turbine cleanup devices. For many of the gas turbine and diesel applications emission standards do not exist. Our goal is to develop coal-fueled diesels and gas turbines that not only meet all applicable emission standards that do exist, but also are capable of meeting possible future standards.

  6. Engineering Tribology

    NASA Astrophysics Data System (ADS)

    Williams, John

    An ideal textbook for a first tribology course, this book provides an interdisciplinary understanding of the field. It includes materials constraints, real design problems and solutions (such as those for journal and rolling element bearing), cams and followers and heavily loaded gear teeth. Including physics, materials science, and surface and lubricant chemistry, the volume integrates quantitative material from a wide variety of disciplines with traditional engineering approaches.

  7. Web Engineering

    SciTech Connect

    White, Bebo

    2003-06-23

    Web Engineering is the application of systematic, disciplined and quantifiable approaches to development, operation, and maintenance of Web-based applications. It is both a pro-active approach and a growing collection of theoretical and empirical research in Web application development. This paper gives an overview of Web Engineering by addressing the questions: (a) why is it needed? (b) what is its domain of operation? (c) how does it help and what should it do to improve Web application development? and (d) how should it be incorporated in education and training? The paper discusses the significant differences that exist between Web applications and conventional software, the taxonomy of Web applications, the progress made so far and the research issues and experience of creating a specialization at the master's level. The paper reaches a conclusion that Web Engineering at this stage is a moving target since Web technologies are constantly evolving, making new types of applications possible, which in turn may require innovations in how they are built, deployed and maintained.

  8. Engineering Liver

    PubMed Central

    Griffith, Linda G.; Wells, Alan; Stolz, Donna Beer

    2014-01-01

    Interest in “engineering liver” arises from multiple communities: therapeutic replacement; mechanistic models of human processes; and drug safety and efficacy studies. An explosion of micro- and nano-fabrication, biomaterials, microfluidic, and other technologies potentially afford unprecedented opportunity to create microphysiological models of human liver, but engineering design principles for how to deploy these tools effectively towards specific applications, including how to define the essential constraints of any given application (including available sources of cells, acceptable cost, and user-friendliness) are still emerging. Arguably less appreciated is the parallel growth in computational systems biology approaches towards these same problems – particularly, in parsing complex disease processes from clinical material, building models of response networks, and in how to interpret the growing compendium of data on drug efficacy and toxicology in patient populations. Here, we provide insight into how the complementary paths of “engineering liver” – experimental and computational – are beginning to interplay towards greater illumination of human disease states and technologies for drug development. PMID:24668880

  9. Planetary engineering

    NASA Technical Reports Server (NTRS)

    Pollack, James B.; Sagan, Carl

    1991-01-01

    Assuming commercial fusion power, heavy lift vehicles and major advances in genetic engineering, the authors survey possible late-21st century methods of working major transformations in planetary environments. Much more Earthlike climates may be produced on Mars by generating low freezing point greenhouse gases from indigenous materials; on Venus by biological conversion of CO2 to graphite, by canceling the greenhouse effect with high-altitude absorbing fine particles, or by a sunshield at the first Lagrangian point; and on Titan by greenhouses and/or fusion warming. However, in our present state of ignorance we cannot guarantee a stable endstate or exclude unanticipated climatic feedbacks or other unintended consequences. Moreover, as the authors illustrate by several examples, many conceivable modes of planetary engineering are so wasteful of scarce solar system resources and so destructive of important scientific information as to raise profound ethical issues, even if they were economically feasible, which they are not. Global warming on Earth may lead to calls for mitigation by planetary engineering, e.g., emplacement and replenishment of anti-greenhouse layers at high altitudes, or sunshields in space. But here especially we must be concerned about precision, stability, and inadvertent side-effects. The safest and most cost-effective means of countering global warming - beyond, e.g., improved energy efficiency, CFC bans and alternative energy sources - is the continuing reforestation of approximately 2 times 107 sq km of the Earth's surface. This can be accomplished with present technology and probably at the least cost.

  10. Engineering Design Handbook: Value Engineering

    DTIC Science & Technology

    1971-07-20

    worthwhile contribution to those goings-on, is often as important to the employee as the amount of compensation he receives. Training in the elements...be set up for: top management, operational management, operating personnel, and value engineers. 4-2.3.1 TOP Managiement Briefings In these...that is what you are being paid to handle, and the level of compensation is very likely to be (or should be) commensurate with your efficacy in

  11. Mission engineering

    NASA Technical Reports Server (NTRS)

    Ondrus, Paul; Fatig, Michael

    1993-01-01

    Goddard Space Flight Center's projects are facing new challenges with respect to the cost effective development and operation of spaceflight missions. Challenges, such as cost limits, compression of schedules, rapidly changing technology, and increasing mission complexity are making the mission development process more dynamic. A concept of 'Mission Engineering' as a means of addressing these challenges is proposed. It is an end-to-end, multimission development methodology that seeks to integrate the development processes between the space, ground, science, and operations segments of a mission. It thereby promotes more mission-oriented system solutions, within and across missions.

  12. Software engineering as an engineering discipline

    NASA Technical Reports Server (NTRS)

    Gibbs, Norman

    1988-01-01

    The goals of the Software Engineering Institute's Education Program are as follows: to increase the number of highly qualified software engineers--new software engineers and existing practitioners; and to be the leading center of expertise for software engineering education and training. A discussion of these goals is presented in vugraph form.

  13. Rotary engine

    SciTech Connect

    Brownfield, L.A.

    1980-12-02

    The major components of this rotary engine are two equal sized rotary units, the housing containing them along with associated ignition and cooling systems. Each of the rotary units consists of a shaft, gear, two outer compressor wheels, and one center power wheel which has twice the axial thickness as the compressor wheel. All the wheels are cylindrical in shape with a lobe section comprising a 180/sup 0/ arc on the periphery of each wheel which forms an expanding and contracting volumetric chamber by means of leading and trailing lips. The lobes of the first rotary unit are situated 180/sup 0/ opposite the lobes of the second adjacent mating rotary unit, thus lobes can intermesh with its corresponding wheel.

  14. Radial engine

    SciTech Connect

    Kmicikiewicz, M.A.

    1988-03-01

    A radial engine is described comprising: a housing; equally spaced openings disposed in ring-like arrangement on the periphery of the housing; a piston and cylinder arrangement in each of the opening, a piston rod for each arrangement fixed to and extending radially inwardly from its respective piston and through its respective opening; shoe means pivotally attached at the other end of each of the piston rod; radial guide means extending in the housing in line with each of the piston rods, and the shoe means provided with guide means followers to ensure radial reciprocal movement of the piston rods and shoe means; and a connecting ring journaled on a crankshaft for circular translation motion in the housing, the ring including a circular rim. Each shoe means includes an arcuate follower member being slidably connected to the rim of the connecting ring.

  15. How Engineers Engineer: Lessons from My First Big Engineering Project

    ERIC Educational Resources Information Center

    Roman, Harry T.

    2008-01-01

    Little did the author realize how much his first engineering project would change his career path, but when it came, he was hooked forever on doing R&D-type engineering. In this article, the author takes the reader back to his first really important electrical engineering project. While the technology he worked on back then is antiquated by…

  16. The responsibilities of engineers.

    PubMed

    Smith, Justin; Gardoni, Paolo; Murphy, Colleen

    2014-06-01

    Knowledge of the responsibilities of engineers is the foundation for answering ethical questions about the work of engineers. This paper defines the responsibilities of engineers by considering what constitutes the nature of engineering as a particular form of activity. Specifically, this paper focuses on the ethical responsibilities of engineers qua engineers. Such responsibilities refer to the duties acquired in virtue of being a member of a group. We examine the practice of engineering, drawing on the idea of practices developed by philosopher Alasdair MacIntyre, and show how the idea of a practice is important for identifying and justifying the responsibilities of engineers. To demonstrate the contribution that knowledge of the responsibilities of engineers makes to engineering ethics, a case study from structural engineering is discussed. The discussion of the failure of the Sleipner A Platform off the coast of Norway in 1991 demonstrates how the responsibilities of engineers can be derived from knowledge of the nature of engineering and its context.

  17. Industrial Education. "Small Engines".

    ERIC Educational Resources Information Center

    Parma City School District, OH.

    Part of a series of curriculum guides dealing with industrial education in junior high schools, this guide provides the student with information and manipulative experiences on small gasoline engines. Included are sections on shop adjustment, safety, small engines, internal combustion, engine construction, four stroke engines, two stroke engines,…

  18. Metabolic Engineering X Conference

    SciTech Connect

    Flach, Evan

    2015-05-07

    The International Metabolic Engineering Society (IMES) and the Society for Biological Engineering (SBE), both technological communities of the American Institute of Chemical Engineers (AIChE), hosted the Metabolic Engineering X Conference (ME-X) on June 15-19, 2014 at the Westin Bayshore in Vancouver, British Columbia. It attracted 395 metabolic engineers from academia, industry and government from around the globe.

  19. A Collaborative Analysis Tool for Integrated Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy; Alexander, Reginald; Landrum, Brian

    2000-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. Then there is a transition to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scramjet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance. To adequately determine the performance of the engine/vehicle, the Hypersonic Flight Inlet Model (HYFIM) module was designed to interface with the RBCC

  20. Engine roughness control means

    SciTech Connect

    Matsuura, M.; Doi, N.; Yoshioka, S.; Okimoto, H.; Veda, K.

    1987-08-04

    This patent describes a control system for a vehicle engine comprising engine condition detecting means for detecting an engine operating condition and producing an engine condition signal representing the engine operating condition, engine combustion control means for controlling a condition of combustion in the engine; and a control factor storage means for storing control factors for controlling the engine combustion. A modifying means connect the comparator means to receive the output signal and to modify the control factor from the storage means by the output of the comparator means so that the combustion control means is controlled by the modified control factor in a direction that the engine vibrations are suppressed. A reference signal changes means connected with the engine condition detecting means to change the reference roughness signal in accordance with the engine operating condition so that the reference signal is decreased when the engine is in idling operation.

  1. Annular nozzle engine technology

    NASA Technical Reports Server (NTRS)

    Martinez, AL

    1992-01-01

    The topics covered include: (1) driver rocket subsystem; (2) annular nozzle engine technology; (3) expansion-deflection nozzle; (4) aerospike-nozzled engine background; (5) aerospike testing; (6) linear aerospike; and (7) the combined cycle engine.

  2. Rotary engine

    SciTech Connect

    Russell, R.L.

    1987-03-31

    An internal combustion four cycle rotary engine is described comprising: a generally cylindrical having one or more accurately spaced cylinders, each carrying a piston therein extending radially of a central rotational axis of the rotor; stationary bearings support shaft means disposed coaxially of the rotor, unitary combustion chamber means carrying main bearing means for rotatably supporting the same on the shaft means and providing one or more individual combustion chambers, each independently communicating with one of the cylinders; the chamber means being mounted concentrically of the rotor and rotatably moveable therewith about the shaft means; cam means comprising a pair of registeringly aligned, axially spaced, continuously curvilinear cam track means which are formed radially assymmetrical about a central axis coincident with the rotational axis of the rotor; the pair of cam track means being located axially outwardly of the cylinders in parallel planes lying formal to the rotational axis and adjacent opposite axial ends of the rotor; cam rider assembly means, each having follower means engaged with the track means for following the contour thereof; and means coupling a rider assembly means to the piston in each cylinder whereby to effect reciprocal strokes of each piston coaxially of its associated cylinder and radially of the rotor in response to the movements of the follower means along the track means; the track means being constructed and arranged to produce distinctly dissimilar movements of the pistons, to produce strokes of unequal duration and length during the respective intake, compression, combustion and exhaust strokes thereof.

  3. Engineering allostery

    DOE PAGES

    Raman, Srivatsan; Taylor, Noah; Genuth, Naomi; ...

    2014-10-08

    Allosteric proteins have great potential in synthetic biology, but our limited understanding of the molecular underpinnings of allostery has hindered the development of designer molecules, including transcription factors with new DNA-binding or ligand-binding specificities that respond appropriately to inducers. Such allosteric proteins could function as novel switches in complex circuits, metabolite sensors, or as orthogonal regulators for independent, inducible control of multiple genes. Advances in DNA synthesis and next-generation sequencing technologies have enabled the assessment of millions of mutants in a single experiment, providing new opportunities to study allostery. Using the classic LacI protein as an example, in this papermore » we describe a genetic selection system using a bidirectional reporter to capture mutants in both allosteric states, allowing the positions most crucial for allostery to be identified. Finally, this approach is not limited to bacterial transcription factors, and could reveal new mechanistic insights and facilitate engineering of other major classes of allosteric proteins such as nuclear receptors, two-component systems, G protein-coupled receptors, and protein kinases.« less

  4. Engineering allostery

    SciTech Connect

    Raman, Srivatsan; Taylor, Noah; Genuth, Naomi; Fields, Stanley; Church, George M.

    2014-10-08

    Allosteric proteins have great potential in synthetic biology, but our limited understanding of the molecular underpinnings of allostery has hindered the development of designer molecules, including transcription factors with new DNA-binding or ligand-binding specificities that respond appropriately to inducers. Such allosteric proteins could function as novel switches in complex circuits, metabolite sensors, or as orthogonal regulators for independent, inducible control of multiple genes. Advances in DNA synthesis and next-generation sequencing technologies have enabled the assessment of millions of mutants in a single experiment, providing new opportunities to study allostery. Using the classic LacI protein as an example, in this paper we describe a genetic selection system using a bidirectional reporter to capture mutants in both allosteric states, allowing the positions most crucial for allostery to be identified. Finally, this approach is not limited to bacterial transcription factors, and could reveal new mechanistic insights and facilitate engineering of other major classes of allosteric proteins such as nuclear receptors, two-component systems, G protein-coupled receptors, and protein kinases.

  5. Rotary engine

    SciTech Connect

    Wilson, Z.

    1990-08-28

    This paper discusses an engine. It comprises a cylinder block; cylinders contained in the cylinder block; matching pistons, each piston reciprocatingly received in one of the cylinders; matching piston rods, each rod connected to one of the pistons and extending outwardly from the block; sheave gears, each sheave gear having a sheave gear axis and a circumference disposed about the sheave gear axis bearing a set of gear teeth thereon; means connecting a respective one of the sheave gears to a respective one of the pistons rods such that reciprocation of the pistons in the cylinders causes rotation of the sheave gears about the sheave gears axes; a combination flywheel/ring gear having a ring gear axis and an outer circumference disposed about the axis bearing a set of ring gear teeth thereon; and means positioning the flywheel/ring gear such that the gear teeth on the flywheel/ring gear engaged the gear teeth on the sheave gears. The flywheel/ring gear is rotated about by its axis by rotation of the sheave gears upon reciprocation of the pistons in the cylinders.

  6. Taking an engine`s temperature

    SciTech Connect

    Allison, S.W.; Beshears, D.L.; Cates, M.R.; Noel, B.W.; Turley, W.D.

    1997-01-01

    Ceramic and ceramic-coated components will be of increasing importance in the advanced engines now under development. Ceramics enable engines to run at much higher temperatures than the superalloys in more conventional engines can. The two options for noncontact high-temperature measurements of ceramic components are pyrometry and phosphor thermometry. This article describes how when properly applied as a thin coating, thermally sensitive phosphors can monitor the temperature of ceramic surfaces inside an engine.

  7. Mod II engine development

    NASA Technical Reports Server (NTRS)

    Karl, David W.

    1987-01-01

    The Mod II engine, a four-cylinder, automotive Stirling engine utilizing the Siemens-Rinia double-acting concept, was assembled and became operational in January 1986. This paper describes the Mod II engine, its first assembly, and the subsequent development work done on engine components up to the point that engine performance characterization testing took place. Performance data for the engine are included.

  8. An Engineering Context for Software Engineering

    DTIC Science & Technology

    2008-09-01

    from a database. Richard Fairley presents one of the definitions reminiscent of industrial engineering practice, ... the technological and...2002. [Fair85] Fairley , Richard, Software Engineering Concepts, McGraw-Hill, 1985. [Flo87] Florman, Samuel C., The Civilized Engineer, St. Martins

  9. Software engineering as an engineering discipline

    NASA Technical Reports Server (NTRS)

    Freedman, Glenn B.

    1988-01-01

    The purpose of this panel is to explore the emerging field of software engineering from a variety of perspectives: university programs; industry training and definition; government development; and technology transfer. In doing this, the panel will address the issues of distinctions among software engineering, computer science, and computer hardware engineering as they relate to the challenges of large, complex systems.

  10. Engineering Encounters: Blasting off with Engineering

    ERIC Educational Resources Information Center

    Dare, Emily A.; Childs, Gregory T.; Cannaday, E. Ashley; Roehrig, Gillian H

    2014-01-01

    What better way to engage young students in physical science concepts than to have them engineer flying toy rockets? The integration of engineering into science classrooms is advocated by the "Next Generation Science Standards" (NGSS) and researchers alike (Brophy et al. 2008), as engineering provides: (1) A "real-world…

  11. Environmental Engineering in Mining Engineering Education

    ERIC Educational Resources Information Center

    Mahamud-Lopez, Manuel Maria; Menendez-Aguado, Juan Maria

    2005-01-01

    In this paper, the current profile of the environmental engineer and the programming of the subject "Environmental Engineering and Technology" corresponding to the studies of Mining Engineering at the University of Oviedo in Spain, is discussed. Professional profile, student knowledge prior to and following instruction as well as…

  12. Space engine safety system

    NASA Technical Reports Server (NTRS)

    Maul, William A.; Meyer, Claudia M.

    1991-01-01

    A rocket engine safety system is designed to initiate control procedures which will minimize damage to the engine and vehicle or test stand in the event of an engine failure. This report describes the features and the implementation issues associated with rocket engine safety systems. Specific concerns of safety systems applied to a space-based engine and long duration space missions are discussed. Examples of safety system features and architectures are given from recent safety monitoring investigations conducted for the Space Shuttle Main Engine and for future liquid rocket engines. Also, a general design and implementation process for rocket engine safety systems is presented.

  13. Subscale Carbon-Carbon Nozzle Extension Development and Hot Fire Testing in Support of Upper Stage Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Gradl, Paul; Valentine, Peter; Crisanti, Matthew; Greene, Sandy Elam

    2016-01-01

    Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures increasing exhaust velocities. Due to the large size of such nozzles and the related engine performance requirements, carbon-carbon (C/C) composite nozzle extensions are being considered for use in order to reduce weight impacts. NASA and industry partner Carbon-Carbon Advanced Technologies (C-CAT) are working towards advancing the technology readiness level of large-scale, domestically-fabricated, C/C nozzle extensions. These C/C extensions have the ability to reduce the overall costs of extensions relative to heritage metallic and composite extensions and to decrease weight by 50%. Material process and coating developments have advanced over the last several years, but hot fire testing to fully evaluate C/C nozzle extensions in relevant environments has been very limited. NASA and C-CAT have designed, fabricated and hot fire tested multiple subscale nozzle extension test articles of various C/C material systems, with the goal of assessing and advancing the manufacturability of these domestically producible materials as well as characterizing their performance when subjected to the typical environments found in a variety of liquid rocket and scramjet engines. Testing at the MSFC Test Stand 115 evaluated heritage and state-of-the-art C/C materials and coatings, demonstrating the capabilities of the high temperature materials and their fabrication methods. This paper discusses the design and fabrication of the 1.2k-lbf sized carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.

  14. Genome scale engineering techniques for metabolic engineering.

    PubMed

    Liu, Rongming; Bassalo, Marcelo C; Zeitoun, Ramsey I; Gill, Ryan T

    2015-11-01

    Metabolic engineering has expanded from a focus on designs requiring a small number of genetic modifications to increasingly complex designs driven by advances in genome-scale engineering technologies. Metabolic engineering has been generally defined by the use of iterative cycles of rational genome modifications, strain analysis and characterization, and a synthesis step that fuels additional hypothesis generation. This cycle mirrors the Design-Build-Test-Learn cycle followed throughout various engineering fields that has recently become a defining aspect of synthetic biology. This review will attempt to summarize recent genome-scale design, build, test, and learn technologies and relate their use to a range of metabolic engineering applications.

  15. Using Collaborative Engineering to Inform Collaboration Engineering

    NASA Technical Reports Server (NTRS)

    Cooper, Lynne P.

    2012-01-01

    Collaboration is a critical competency for modern organizations as they struggle to compete in an increasingly complex, global environment. A large body of research on collaboration in the workplace focuses both on teams, investigating how groups use teamwork to perform their task work, and on the use of information systems to support team processes ("collaboration engineering"). This research essay presents collaboration from an engineering perspective ("collaborative engineering"). It uses examples from professional and student engineering teams to illustrate key differences in collaborative versus collaboration engineering and investigates how challenges in the former can inform opportunities for the latter.

  16. Alternative Automobile Engines

    ERIC Educational Resources Information Center

    Wilson, David Gordon

    1978-01-01

    Requirements for cleaner and more efficient engines have stimulated a search for alternatives to the conventional spark-ignition engine. So far, the defects of the alternative engines are clearer than the virtues. The following engines are compared: spark ignition, diesel, vapor-cycle, Stirling, and gas turbine. (Author/MA)

  17. Structural design and analysis of a Mach zero to five turbo-ramjet system

    NASA Technical Reports Server (NTRS)

    Spoth, Kevin A.; Moses, Paul L.

    1993-01-01

    The paper discusses the structural design and analysis of a Mach zero to five turbo-ramjet propulsion system for a Mach five waverider-derived cruise vehicle. The level of analysis detail necessary for a credible conceptual design is shown. The results of a finite-element failure mode sizing analysis for the engine primary structure is presented. The importance of engine/airframe integration is also discussed.

  18. Exo-Skeletal Engine: Novel Engine Concept

    NASA Technical Reports Server (NTRS)

    Chamis, Cristos C.; Blankson, Isaiah M.

    2004-01-01

    The exo-skeletal engine concept represents a new radical engine technology with the potential to substantially revolutionize engine design. It is an all-composite drum-rotor engine in which conventionally heavy shafts and discs are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Thus the rotating blades are in compression rather than tension. The resulting open channel at the engine centerline has immense potential for jet noise reduction and can also accommodate an inner combined-cycle thruster such as a ramjet. The exo-skeletal engine is described in some detail with respect to geometry, components, and potential benefits. Initial evaluations and results for drum rotors, bearings, and weights are summarized. Component configuration, assembly plan, and potential fabrication processes are also identified. A finite element model of the assembled engine and its major components is described. Preliminary results obtained thus far show at least a 30-percent reduction of engine weight and about a 10-dB noise reduction, compared with a baseline conventional high-bypass-ratio engine. Potential benefits in all aspects of this engine technology are identified and tabulated. Quantitative assessments of potential benefits are in progress.

  19. Engineering Lessons Learned and Systems Engineering Applications

    NASA Technical Reports Server (NTRS)

    Gill, Paul S.; Garcia, Danny; Vaughan, William W.

    2005-01-01

    Systems Engineering is fundamental to good engineering, which in turn depends on the integration and application of engineering lessons learned and technical standards. Thus, good Systems Engineering also depends on systems engineering lessons learned from within the aerospace industry being documented and applied. About ten percent of the engineering lessons learned documented in the NASA Lessons Learned Information System are directly related to Systems Engineering. A key issue associated with lessons learned datasets is the communication and incorporation of this information into engineering processes. Systems Engineering has been defined (EINIS-632) as "an interdisciplinary approach encompassing the entire technical effort to evolve and verify an integrated and life-cycle balanced set of system people, product, and process solutions that satisfy customer needs". Designing reliable space-based systems has always been a goal for NASA, and many painful lessons have been learned along the way. One of the continuing functions of a system engineer is to compile development and operations "lessons learned" documents and ensure their integration into future systems development activities. They can produce insights and information for risk identification identification and characterization. on a new project. Lessons learned files from previous projects are especially valuable in risk

  20. Application of CARS to scramjet combustion

    NASA Technical Reports Server (NTRS)

    Antcliff, R. R.

    1987-01-01

    A coherent anti-Stokes Raman spectroscopic (CARS) instrument has been developed for measuring simultaneously temperature and N2 - O2 species concentration in hostile flame environments. A folded BOXCARS arrangement was employed to obtain high spatial resolution. Polarization discrimination against the nonresonant background decreased the lower limits of O2 detectivity. The instrument has been primarily employed for validation of computational fluid-dynamics computer-model codes. Comparisons have been made to both the CHARNAL and TEACH codes on a hydrogen diffusion flame with good results.

  1. Shock tunnel studies of scramjet phenomena 1994

    NASA Technical Reports Server (NTRS)

    1995-01-01

    A new expansion tube facility has been built, and is in the process of being commissioned. It has a bore of 90 mm, and has been designed for peak rupture pressures of 100 Mpa. It is configured with multiple sections and diaphragm location stations to give optimized performance over a range of sub and superorbital conditions. It has a compound piston arrangement for a two stage compression, designed to maximize the length of shock expansion tube which can be driven within a fixed total facility length. Experiments have been successfully performed with a dummy first stage piston, and a rubber energy absorbing brake. The results agree well with a one-dimensional stress wave model of the piston impacting on the rubber, and codes for piston motion. Strain of the rubber is restricted to approximately 20 percent at which level no damage is to be expected in the buffer material, and none has been observed, indicating that the mechanism will be fully reuseable.

  2. Applying MHD Results to a Scramjet Vehicle

    DTIC Science & Technology

    2007-02-12

    1194, Reno, NV, January 2005. 10. Rosa, R.J., "Magnetohydrodynamic Energy Conversion", Hemisphere Publishing Corp., New York, 1968, Revised Printing ...Introduction and Background Projects Analytical and experimental studies of both electrohydrodynamic (EHD) and magnetohydrodynamic (MHD) application to...Institute, Tullahoma, TN, June 1981. 26. Vendell, E.W., "Free- Jet Electrical Conductivity Profiles of a Seeded MHD Combustion Plasma", 15"’ Symposium

  3. Magnetogasdynamic Flow Acceleration in a Scramjet Nozzle

    DTIC Science & Technology

    2004-06-01

    Col. Maple and Lt. Col. Hughson for teaching me the basics of CFD and helping me to appreciate a part of this project that I became so involved in...is large. Sa Assumption 3: Thermochemistry of the flow must be known. Sa Assumption 4: The velocity at each stage is aligned with the thrust or axial

  4. Preliminary calibration of a generic scramjet combustor

    NASA Technical Reports Server (NTRS)

    Jacobs, P. A.; Morgan, R. G.; Rogers, R. C.; Wendt, M.; Brescianini, C.; Paull, A.; Kelly, G.

    1991-01-01

    The results of a preliminary investigation of the combustion of hydrogen fuel at hypersonic flow conditions are provided. The tests were performed in a generic, constant-area combustor model with test gas supplied by a free-piston-driven reflected-shock tunnel. Static pressure measurements along the combustor wall indicated that burning did occur for combustor inlet conditions of P(static) approximately equal to 19kPa, T(static) approximately equal to 1080 K, and U approximately equal to 3630 m/s with a fuel equivalence ratio approximately equal to 0.9. These inlet conditions were obtained by operating the tunnel with stagnation enthalpy approximately equal to 8.1 MJ/kg, stagnation pressure approximately equal to 52 MPa, and a contoured nozzle with a nominal exit Mach number of 5.5.

  5. Heat conduction controlled combustion for scramjet applications

    NASA Technical Reports Server (NTRS)

    Ferri, A.; Agnone, A. M.

    1974-01-01

    The use of heat conduction flame generated in a premixed supersonic stream is discussed. It is shown that the flame is controlled initially by heat conduction and then by chemical reaction. Such a flame is shorter than the diffusion type of flame and therefore it requires a much shorter burner. The mixing is obtained by injecting the hydrogen in the inlet. Then the inlet can be cooled by film cooling.

  6. Shock tunnel studies of scramjet phenomena

    NASA Technical Reports Server (NTRS)

    Morgan, R. G.; Paull, A.; Stalker, R. J.; Jacobs, P.; Morris, N.; Stringer, I.; Brescianini, C.

    1988-01-01

    Commissioning of the new T4 shock tunnel at the University of Queensland implied that it was no longer necessary to focus the work of the research group about an annual test series conducted in the T3 shock tunnel in Canberra. Therefore, it has been possible to organize a group for work to proceed along lines such that particular personnel are associated with particular project areas. The format of this report consists of a series of reports on specific project areas, with a brief general introduction commenting on each report. The introduction is structured by project areas, with the title of the relevant report stated under the project area heading. The reports themselves follow in the order of the project area headings.

  7. Scramjet Flow Field Control Using Magnetogasdynamics

    DTIC Science & Technology

    2006-04-01

    2-9 ηKE Kinetic energy efficiency . . . . . . . . . . . . . . . . . . . 2-9 ∆s/Cp Dimensionless entropy increase...one, Q = σ0B 2 0L0/(ρ0V0). These assumptions lead to a simplified combination of the Maxwell equations for elec- tromagnetics and the Navier-Stokes...energy efficiency, ηKE, the dimensionless entropy increase, ∆s/Cp, and the adiabatic compression efficiency, ηc [15]. πc and ∆s/Cp both rely on constant

  8. Diode Laser Sensor for Scramjet Inlets

    DTIC Science & Technology

    2011-06-03

    Capacitor SM 0805 (2012) 1 C48, C49, C50 , C52 10p Capacitor SM 0805 (2012) 4 C62, C64, C66 10uF Capacitor TANT B (3528) 3 D1, D2 BAT54 Schottky Diode...ApprovedOMB No. 0704-0188 Public reporting burden for the collection of information is estimated to average 1 hour per response, including the time for...with a collection of information if it does not display a currently valid OMB control number. 1 . REPORT DATE 07 JUN 2011 2. REPORT TYPE Final 3

  9. Diode Laser Sensor for Scramjet Inlet

    DTIC Science & Technology

    2010-05-11

    Conference’. 1.2 O’Byrne, S., Huynh, L., Wittig, S. M. and Smith, N. S. A. (2009), Non- intrusive water vapour absorp- tion measurements in a simulated...O’Byrne, L. Huynh, S. M. Wittig and N. S. A. Smith, “Non- intrusive Water Vapour Absorp- tion Measurements in a Simulated Helicopter Exhaust”, Proceedings...rather than at a surface. The measurement techniques used at these hypersonic flow conditions should also be non- intrusive . Tuneable diode laser

  10. Software Engineering Guidebook

    NASA Technical Reports Server (NTRS)

    Connell, John; Wenneson, Greg

    1993-01-01

    The Software Engineering Guidebook describes SEPG (Software Engineering Process Group) supported processes and techniques for engineering quality software in NASA environments. Three process models are supported: structured, object-oriented, and evolutionary rapid-prototyping. The guidebook covers software life-cycles, engineering, assurance, and configuration management. The guidebook is written for managers and engineers who manage, develop, enhance, and/or maintain software under the Computer Software Services Contract.

  11. Inlet Engineering Toolbox

    DTIC Science & Technology

    2014-10-31

    ADDRESS(ES) U.S. Army Engineer Research and Development Center,CIRP - The Coastal Inlets Research Program,3909 Halls Ferry Road,Vicksburg,MS,39180... Coastal Inlets Research Program Inlet Engineering Toolbox The Inlet Engineering Toolbox (IET) Work Unit develops desktop PC and web-based tools to...aid in studies of the consequences of engineering actions at coastal inlets and adjacent beaches. District scientists and engineers need rapid

  12. Biomedical engineering education through global engineering teams.

    PubMed

    Scheffer, C; Blanckenberg, M; Garth-Davis, B; Eisenberg, M

    2012-01-01

    Most industrial projects require a team of engineers from a variety of disciplines. The team members are often culturally diverse and geographically dispersed. Many students do not acquire sufficient skills from typical university courses to function efficiently in such an environment. The Global Engineering Teams (GET) programme was designed to prepare students such a scenario in industry. This paper discusses five biomedical engineering themed projects completed by GET students. The benefits and success of the programme in educating students in the field of biomedical engineering are discussed.

  13. Teaching Engineering Practices

    NASA Astrophysics Data System (ADS)

    Cunningham, Christine M.; Carlsen, William S.

    2014-03-01

    Engineering is featured prominently in the Next Generation Science Standards (NGSS) and related reform documents, but how its nature and methods are described is problematic. This paper is a systematic review and critique of that representation, and proposes that the disciplinary core ideas of engineering (as described in the NGSS) can be disregarded safely if the practices of engineering are better articulated and modeled through student engagement in engineering projects. A clearer distinction between science and engineering practices is outlined, and prior research is described that suggests that precollege engineering design can strengthen children's understandings about scientific concepts. However, a piecemeal approach to teaching engineering practices is unlikely to result in students understanding engineering as a discipline. The implications for science teacher education are supplemented with lessons learned from a number of engineering education professional development projects.

  14. Humanitarian engineering in the engineering curriculum

    NASA Astrophysics Data System (ADS)

    Vandersteen, Jonathan Daniel James

    There are many opportunities to use engineering skills to improve the conditions for marginalized communities, but our current engineering education praxis does not instruct on how engineering can be a force for human development. In a time of great inequality and exploitation, the desire to work with the impoverished is prevalent, and it has been proposed to adjust the engineering curriculum to include a larger focus on human needs. This proposed curriculum philosophy is called humanitarian engineering. Professional engineers have played an important role in the modern history of power, wealth, economic development, war, and industrialization; they have also contributed to infrastructure, sanitation, and energy sources necessary to meet human need. Engineers are currently at an important point in time when they must look back on their history in order to be more clear about how to move forward. The changing role of the engineer in history puts into context the call for a more balanced, community-centred engineering curriculum. Qualitative, phenomenographic research was conducted in order to understand the need, opportunity, benefits, and limitations of a proposed humanitarian engineering curriculum. The potential role of the engineer in marginalized communities and details regarding what a humanitarian engineering program could look like were also investigated. Thirty-two semi-structured research interviews were conducted in Canada and Ghana in order to collect a pool of understanding before a phenomenographic analysis resulted in five distinct outcome spaces. The data suggests that an effective curriculum design will include teaching technical skills in conjunction with instructing about issues of social justice, social location, cultural awareness, root causes of marginalization, a broader understanding of technology, and unlearning many elements about the role of the engineer and the dominant economic/political ideology. Cross-cultural engineering development

  15. Service Cart For Engines

    NASA Technical Reports Server (NTRS)

    Ng, Gim Shek

    1995-01-01

    Cart supports rear-mounted air-cooled engine from Volkswagen or Porsche automobile. One person removes, repairs, tests, and reinstalls engine of car, van, or home-built airplane. Consists of framework of wood, steel, and aluminum components supported by four wheels. Engine lifted from vehicle by hydraulic jack and gently lowered onto waiting cart. Jack removed from under engine. Rear of vehicle raised just enough that engine can be rolled out from under it. Cart easily supports 200-lb engine. Also used to hold transmission. With removable sheet-metal top, cart used as portable seat.

  16. The Phillips Stirling engine

    SciTech Connect

    Hargreaves, C.M.

    1991-01-01

    This book is about the Stirling engine and its development from the heavy cast-iron machine of the 19th century to that of today. It is a history of a research effort spanning nearly 50 years, together with an outline of principles, and some technical details and descriptions of the more important engines. Contents include: the hot-air engine; the 20th-century revival; the Stirling cycle; rhombic-drive engines; heating and cooling; pistons and seals; electric generators and heat pumps; exotic heat sources; the engine and the environment; swashplate engines; and the past and the future.

  17. Improving engineering effectiveness

    NASA Technical Reports Server (NTRS)

    Fiero, J. D.

    1985-01-01

    Methodologies to improve engineering productivity were investigated. The rocky road to improving engineering effectiveness is reviewed utilizing a specific semiconductor engineering organization as a case study. The organization had a performance problem regarding new product introductions. With the help of this consultant as a change agent the engineering team used a systems approach to through variables that were effecting their output significantly. Critical factors for improving this engineering organization's effectiveness and the roles/responsibilities of management, the individual engineers and the internal consultant are discussed.

  18. Tripropellant engine study

    NASA Technical Reports Server (NTRS)

    Wheeler, D. B.

    1978-01-01

    Engine performance data, combustion gas thermodynamic properties, and turbine gas parameters were determined for various high power cycle engine configurations derived from the space shuttle main engine that will allow sequential burning of LOX/hydrocarbon and LOX/hydrogen fuels. Both stage combustion and gas generator pump power cycles were considered. Engine concepts were formulated for LOX/RP-1, LOX/CH4, and LOX/C3H8 propellants. Flowrates and operating conditions were established for this initial set of engine systems, and the adaptability of the major components of shuttle main engine was investigated.

  19. Engine valve driving apparatus

    SciTech Connect

    Masuda, S.; Uesugi, T.; Oda, H.

    1989-01-03

    An engine valve driving apparatus for an internal combustion engine having a cam driven engine valve is described. It consists of a camshaft rotatable in synchronism with rotation of a crankshaft of an engine and a movable cam member supported by the camshaft for axial movement and prevented from turning relative to the camshaft. The movable cam member can be axially shifted between an operative position wherein the cam member is cooperative with a member of the engine valve so as to cause an operation of the engine valve and an inoperative position wherein the cam member is out of cooperation with the member.

  20. Engineering Ethics in the Subject of Engineering History

    NASA Astrophysics Data System (ADS)

    Isohata, Hiroshi

    Engineering ethics has been focused in the field of engineering education since the introduction of accreditation system of engineering education. In this paper, contents of the subject of engineering history are examined and discussed from the viewpoints of education of engineering ethics through a practical case of civil engineering history in a college. For the first step, codes of engineering ethics regulated in various engineering organizations are analyzed and the common contents are extracted to set the requirements for the education of engineering ethics. Then contents of the subject of engineering history are examined according to the requirements. Finally, conditions of engineering history for engineering ethics are discussed.

  1. Probable Impacts of Space Operations on Air Force Civil Engineering

    DTIC Science & Technology

    1989-05-01

    University, Air Command and Staff College, Research Report). Bierling , James R ., Space Operations Professional Development Guide, Maxwell AFB, Alabama, April...basing containerized payloads self ferry Table S. NASP Technical Challenges and Confidence Factors4 1986 1988 1990 Airframe Structures and Materials R ...Y B Thermal Management Y G B Flight Vehicle Integration R G G Inlet/Nozzle Performane R Y 6 Slush Hydrogen Y Y 6 Propulsion Ramjet Y 6 B Scramjet

  2. Engineering and Engineering Technology Degrees, 1990.

    ERIC Educational Resources Information Center

    Ellis, R. A.

    1991-01-01

    The number of B.S., M.S./P.E., and Ph.D. degrees in engineering and engineering technology awarded by U.S. colleges and universities is tabulated according to the following criteria: by state, by school, curriculum, type of recipient, and by school and degree level. (KR)

  3. Advanced Natural Gas Reciprocating Engine(s)

    SciTech Connect

    Pike, Edward

    2014-03-31

    The objective of the Cummins ARES program, in partnership with the US Department of Energy (DOE), is to develop advanced natural gas engine technologies that increase engine system efficiency at lower emissions levels while attaining lower cost of ownership. The goals of the project are to demonstrate engine system achieving 50% Brake Thermal Efficiency (BTE) in three phases, 44%, 47% and 50% (starting baseline efficiency at 36% BTE) and 0.1 g/bhp-hr NOx system out emissions (starting baseline NOx emissions at 2 – 4 g/bhp-hr NOx). Primary path towards above goals include high Brake Mean Effective Pressure (BMEP), improved closed cycle efficiency, increased air handling efficiency and optimized engine subsystems. Cummins has successfully demonstrated each of the phases of this program. All targets have been achieved through application of a combined set of advanced base engine technologies and Waste Heat Recovery from Charge Air and Exhaust streams, optimized and validated on the demonstration engine and other large engines. The following architectures were selected for each Phase: Phase 1: Lean Burn Spark Ignited (SI) Key Technologies: High Efficiency Turbocharging, Higher Efficiency Combustion System. In production on the 60/91L engines. Over 500MW of ARES Phase 1 technology has been sold. Phase 2: Lean Burn Technology with Exhaust Waste Heat Recovery (WHR) System Key Technologies: Advanced Ignition System, Combustion Improvement, Integrated Waste Heat Recovery System. Base engine technologies intended for production within 2 to 3 years Phase 3: Lean Burn Technology with Exhaust and Charge Air Waste Heat Recovery System Key Technologies: Lower Friction, New Cylinder Head Designs, Improved Integrated Waste Heat Recovery System. Intended for production within 5 to 6 years Cummins is committed to the launch of next generation of large advanced NG engines based on ARES technology to be commercialized worldwide.

  4. Twin engine synchronizer

    SciTech Connect

    Kobus, J.R.

    1988-05-03

    This patent describes an apparatus for synchronizing the speeds of two engines, each having its own throttle level connected by an associated cable to a respective hand throttle lever, comprising moving means carried by the throttle lever of one of the engines for moving the throttle lever of the one engine independently of its associated cable and its respective hand throttle lever to increase or decrease the speed of the one engine until the speed of the one engine matches the speed of the other engine. The moving means moves the throttle lever of the one engine without moving its associated cable or its respective hand throttle lever, and actuating means mounted remote from the throttle lever of the one engine for actuating the moving means.

  5. Structural Engineering: Overview

    NASA Technical Reports Server (NTRS)

    Castro, Edgar

    2011-01-01

    This slide presentation presents the work of the Structural Engineering Division of the Engineering Directorate. The work includes: providing technical expertise and leadership for the development, evaluation, and operation of structural, mechanical, and thermal spaceflight systems.

  6. Rotorcraft convertible engine study

    NASA Technical Reports Server (NTRS)

    Gill, J. C.; Earle, R. V.; Mar, H. M.

    1982-01-01

    The objective of the Rotorcraft Convertible Engine Study was to define future research and technology effort required for commercial development by 1988 of convertible fan/shaft gas turbine engines for unconventional rotorcraft transports. Two rotorcraft and their respective missions were defined: a Fold Tilt Rotor aircraft and an Advancing Blade Concept (ABC) rotorcraft. Sensitivity studies were conducted with these rotorcraft to determine parametrically the influence of propulsion characteristics on aircraft size, mission fuel requirements, and direct operating costs (DOC). The two rotorcraft were flown with conventional propulsion systems (separate lift/cruise engines) and with convertible propulsion systems to determine the benefits to be derived from convertible engines. Trade-off studies were conducted to determine the optimum engine cycle and staging arrangement for a convertible engine. Advanced technology options applicable to convertible engines were studied. Research and technology programs were identified which would ensure technology readiness for commercial development of convertible engines by 1988.

  7. NASA systems engineering handbook

    NASA Astrophysics Data System (ADS)

    Shishko, Robert; Aster, Robert; Chamberlain, Robert G.; McDuffee, Patrick; Pieniazek, Les; Rowell, Tom; Bain, Beth; Cox, Renee I.; Mooz, Harold; Polaski, Lou

    1995-06-01

    This handbook brings the fundamental concepts and techniques of systems engineering to NASA personnel in a way that recognizes the nature of NASA systems and environment. It is intended to accompany formal NASA training courses on systems engineering and project management when appropriate, and is designed to be a top-level overview. The concepts were drawn from NASA field center handbooks, NMI's/NHB's, the work of the NASA-wide Systems Engineering Working Group and the Systems Engineering Process Improvement Task team, several non-NASA textbooks and guides, and material from independent systems engineering courses taught to NASA personnel. Five core chapters cover systems engineering fundamentals, the NASA Project Cycle, management issues in systems engineering, systems analysis and modeling, and specialty engineering integration. It is not intended as a directive.

  8. Educating the Engineer.

    ERIC Educational Resources Information Center

    Wallace, Melanie; Wallace, Mack

    2003-01-01

    Presented as a conversation between a teacher and engineer about school design, addresses educators' preferences and engineers' perspectives on issues, such as windows, sustainable design, sinks, acoustics, and natural ventilation. (EV)

  9. Aircraft Engine Emissions. [conference

    NASA Technical Reports Server (NTRS)

    1977-01-01

    A conference on a aircraft engine emissions was held to present the results of recent and current work. Such diverse areas as components, controls, energy efficient engine designs, and noise and pollution reduction are discussed.

  10. The coupling of engines

    NASA Technical Reports Server (NTRS)

    Boccaccio, Paul

    1921-01-01

    This report examines the idea of coupling numerous engines together to turn a single propeller, which the author feels would free aircraft design from the problems of multi-engine and propeller design.

  11. Keeping Engineers Current.

    ERIC Educational Resources Information Center

    Shell, Ellen Ruppel

    1985-01-01

    Notes several options for keeping practicing engineers abreast of the state of the art, including university seminars, videotape courses, and instruction via satellite. Also cites the need for engineers to be encouraged to pursue life-long professional education. (DH)

  12. Chemical Engineering Education Revisited.

    ERIC Educational Resources Information Center

    Theodore, Louis

    1978-01-01

    The opinion is presented that chemical engineering education seems to emphasize the professor's research and/or professional interests with little regard for the real needs of the student who intends to become a practicing engineer. (BB)

  13. NASA Systems Engineering Handbook

    NASA Technical Reports Server (NTRS)

    Shishko, Robert; Aster, Robert; Chamberlain, Robert G.; Mcduffee, Patrick; Pieniazek, Les; Rowell, Tom; Bain, Beth; Cox, Renee I.; Mooz, Harold; Polaski, Lou

    1995-01-01

    This handbook brings the fundamental concepts and techniques of systems engineering to NASA personnel in a way that recognizes the nature of NASA systems and environment. It is intended to accompany formal NASA training courses on systems engineering and project management when appropriate, and is designed to be a top-level overview. The concepts were drawn from NASA field center handbooks, NMI's/NHB's, the work of the NASA-wide Systems Engineering Working Group and the Systems Engineering Process Improvement Task team, several non-NASA textbooks and guides, and material from independent systems engineering courses taught to NASA personnel. Five core chapters cover systems engineering fundamentals, the NASA Project Cycle, management issues in systems engineering, systems analysis and modeling, and specialty engineering integration. It is not intended as a directive. Superseded by: NASA/SP-2007-6105 Rev 1 (20080008301).

  14. Computerized Interactive Harness Engineering

    NASA Technical Reports Server (NTRS)

    Billitti, J. W.

    1985-01-01

    Computerized interactive harness engineering program inexpensive, interactive system for learning and using engineering approach to interconnection systems. Basically data-base system that stores information as files of individual connectors and handles wiring information in circuit groups stored as records.

  15. 40 CFR 90.116 - Certification procedure-determining engine displacement, engine class, and engine families.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... engine displacement, engine class, and engine families. 90.116 Section 90.116 Protection of Environment...-IGNITION ENGINES AT OR BELOW 19 KILOWATTS Emission Standards and Certification Provisions § 90.116 Certification procedure—determining engine displacement, engine class, and engine families. (a)...

  16. 40 CFR 90.116 - Certification procedure-determining engine displacement, engine class, and engine families.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... engine displacement, engine class, and engine families. 90.116 Section 90.116 Protection of Environment...-IGNITION ENGINES AT OR BELOW 19 KILOWATTS Emission Standards and Certification Provisions § 90.116 Certification procedure—determining engine displacement, engine class, and engine families. (a)...

  17. 40 CFR 90.116 - Certification procedure-determining engine displacement, engine class, and engine families.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... engine displacement, engine class, and engine families. 90.116 Section 90.116 Protection of Environment...-IGNITION ENGINES AT OR BELOW 19 KILOWATTS Emission Standards and Certification Provisions § 90.116 Certification procedure—determining engine displacement, engine class, and engine families. (a)...

  18. 40 CFR 90.116 - Certification procedure-determining engine displacement, engine class, and engine families.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... engine displacement, engine class, and engine families. 90.116 Section 90.116 Protection of Environment...-IGNITION ENGINES AT OR BELOW 19 KILOWATTS Emission Standards and Certification Provisions § 90.116 Certification procedure—determining engine displacement, engine class, and engine families. (a)...

  19. Liquid Rocket Engine Testing

    DTIC Science & Technology

    2016-10-21

    Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL...Distribution Unlimited. PA Clearance 16493 Liquid Rocket Engine Testing • Engines and their components are extensively static- tested in development • This

  20. Earthquake engineering research: 1982

    NASA Astrophysics Data System (ADS)

    The Committee on Earthquake Engineering Research addressed two questions: What progress has research produced in earthquake engineering and which elements of the problem should future earthquake engineering pursue. It examined and reported in separate chapters of the report: Applications of Past Research, Assessment of Earthquake Hazard, Earthquake Ground Motion, Soil Mechanics and Earth Structures, Analytical and Experimental Structural Dynamics, Earthquake Design of Structures, Seismic Interaction of Structures and Fluids, Social and Economic Aspects, Earthquake Engineering Education, Research in Japan.

  1. Cruise Missile Engines

    NASA Technical Reports Server (NTRS)

    1982-01-01

    Williams International's F107 fanjet engine is used in two types of cruise missiles, Navy-sponsored Tomahawk and the Air Force AGM-86B Air Launched Cruise Missile (ALCM). Engine produces about 600 pounds thrust, is one foot in diameter and weighs only 141 pounds. Design was aided by use of a COSMIC program in calculating airflows in engine's internal ducting, resulting in a more efficient engine with increased thrust and reduced fuel consumption.

  2. Biomedical Engineering Laboratory

    DTIC Science & Technology

    2007-11-02

    The Masters of Engineering program with concentration in Biomedical Engineering at Tennessee State University was established in fall 2000. Under... biomedical engineering . The lab is fully equipped with 10 Pentium5-based, 2 Pentium4-based laptops for mobile experiments at remote locations, 8 Biopac...students (prospective graduate students in biomedical engineering ) are regularly using this lab. This summer, 8 new prospective graduate students

  3. Deployable Engine Air Brake

    NASA Technical Reports Server (NTRS)

    2014-01-01

    On approach, next-generation aircraft are likely to have airframe noise levels that are comparable to or in excess of engine noise. ATA Engineering, Inc. (ATA) is developing a novel quiet engine air brake (EAB), a device that generates "equivalent drag" within the engine through stream thrust reduction by creating a swirling outflow in the turbofan exhaust nozzle. Two Phase II projects were conducted to mature this technology: (1) a concept development program (CDP) and (2) a system development program (SDP).

  4. Unified Engineering Software System

    NASA Technical Reports Server (NTRS)

    Purves, L. R.; Gordon, S.; Peltzman, A.; Dube, M.

    1989-01-01

    Collection of computer programs performs diverse functions in prototype engineering. NEXUS, NASA Engineering Extendible Unified Software system, is research set of computer programs designed to support full sequence of activities encountered in NASA engineering projects. Sequence spans preliminary design, design analysis, detailed design, manufacturing, assembly, and testing. Primarily addresses process of prototype engineering, task of getting single or small number of copies of product to work. Written in FORTRAN 77 and PROLOG.

  5. COBRA Main Engine Project

    NASA Technical Reports Server (NTRS)

    Snoddy, Jim; Sides, Steve; Lyles, Garry M. (Technical Monitor)

    2002-01-01

    The COBRA (CO-Optimized Booster for Reusable Applications) project include the following: 1. COBRA main engine project team. 2. COBRA and RLX cycles selected. 3. COBRA proto-type engine approach enables mission success. 4. COBRA provides quick, low cost demo of cycle and technologies. 5. COBRA cycle I risk reduction supports. 6. Achieving engine safety. 6. RLX cycle I risk reduction supports. 7. Flight qualification. 9. Life extension engine testing.

  6. Engineering for All

    ERIC Educational Resources Information Center

    Lottero-Perdue, Pamela S.; Lovelidge, Sarah; Bowling, Erin

    2010-01-01

    As calls for science, technology, engineering, and mathematics (STEM) education at the elementary level become more vociferous, elementary teachers may be wondering whether engineering is meant for "all" students. However, the authors assert that engineering can be taught in inclusive environments. It may be especially empowering for those who…

  7. Educating More Engineers

    ERIC Educational Resources Information Center

    Environmental Science and Technology, 1973

    1973-01-01

    Indicates that there will be a substantially increased demand for environmental engineers during the next few years, especially in the areas of water pollution control and sanitary engineering. Educators see the need for additional engineering graduates and for improved environmental training programs in schools. (JR)

  8. Principles of Naval Engineering.

    ERIC Educational Resources Information Center

    Naval Personnel Program Support Activity, Washington, DC.

    Fundamentals of shipboard machinery, equipment, and engineering plants are presented in this text prepared for engineering officers. A general description is included of the development of naval ships, ship design and construction, stability and buoyancy, and damage and casualty control. Engineering theories are explained on the background of ship…

  9. Personality Characteristics of Engineers

    ERIC Educational Resources Information Center

    van der Molen, Henk T.; Schmidt, Henk G.; Kruisman, Gerard

    2007-01-01

    The objective of the current study was to investigate the personality characteristics of a group of engineers with a variety of years of experience. It was executed to remedy shortcomings of the literature concerning this issue and to produce suggestions for a postgraduate training programme for engineers. A total of 103 engineers were tested with…

  10. H-1 Engine

    NASA Technical Reports Server (NTRS)

    1960-01-01

    Alignment of the H-1 engine performed in the Army Ballistic Missile Agency (ABMA ), building 4708, in February 1960. A cluster of eight H-1 engines were used to thrust the first stage of the Saturn I launch vehicle. The H-1 engine was developed under the direction of the Marshall Space Flight Center.

  11. H-1 Engines

    NASA Technical Reports Server (NTRS)

    1959-01-01

    The H-1 engines for the Saturn I vehicle in the alignment fixture. A cluster of eight H-1 engines were used to thrust the first stage of the Saturn I launch vehicle. The H-1 engine was developed under the direction of the Marshall Space Flight Center.

  12. Engineering for Native Americans.

    ERIC Educational Resources Information Center

    Jarosz, Jeffrey

    2003-01-01

    The engineering workforce is overwhelmingly male and White. To attract and retain American Indian and other minority-group students, engineering programs must offer practical, hands-on, team activities; show that engineering is beneficial to society and Indian communities; use inclusive textbooks; offer distance-learning opportunities; and…

  13. The Engineering Manpower Shortage.

    ERIC Educational Resources Information Center

    Keyworth, George A.

    1982-01-01

    The author discusses the shortage of qualified engineers and the importance to the United States of high-quality engineering faculty, high-quality engineering graduates, and strong math and science departments in high schools. The author argues that it is up to the private sector to encourage these trends. (CT)

  14. Program (systems) engineering

    NASA Technical Reports Server (NTRS)

    Baroff, Lynn E.; Easter, Robert W.; Pomphrey, Richard B.

    2004-01-01

    Program Systems Engineering applies the principles of Systems Engineering at the program level. Space programs are composed of interrelated elements which can include collections of projects, advanced technologies, information systems, etc. Some program elements are outside traditional engineering's physical systems, such as education and public outreach, public relations, resource flow, and interactions within the political environments.

  15. 20. Engine identified as a 'single cylinder vacuum assist engine ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    20. Engine identified as a 'single cylinder vacuum assist engine for the Tod tandem compound engine' showing crank end. - Carnegie Steel-Ohio Works, Steam Engines, 912 Salt Springs Road, Youngstown, Mahoning County, OH

  16. 21. Engine identified as a 'single cylinder vacuum assist engine ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    21. Engine identified as a 'single cylinder vacuum assist engine for Tod tandem compound engine' showing compressor. - Carnegie Steel-Ohio Works, Steam Engines, 912 Salt Springs Road, Youngstown, Mahoning County, OH

  17. Overview With Results and Lessons Learned of the X-43A Mach 10 Flight

    NASA Technical Reports Server (NTRS)

    Marshall, Laurie A.; Bahm, Catherine; Corpening, Griffin P.; Sherrill, Robert

    2005-01-01

    This paper provides an overview of the final flight of the NASA X-43A project. The project consisted of three flights, two planned for Mach 7 and one for Mach 10. The third and final flight, November 16, 2004, was the first Mach 10 flight demonstration of an airframe-integrated, scramjet-powered, hypersonic vehicle. The goals and objectives for the project as well as those for the third flight are presented. The configuration of the Hyper-X stack including the X-43A, Hyper-X launch vehicle, and Hyper-X research vehicle adapter is discussed. The second flight of the X-43A was successfully conducted on March 27, 2004. Mission differences, vehicle modifications and lessons learned from the second flight as they applied to the third flight are also discussed. An overview of flight 3 results is presented.

  18. Hyper-X Program Status

    NASA Technical Reports Server (NTRS)

    McClinton, Charles R.; Reubush, David E.; Sitz, Joel; Reukauf, Paul

    2001-01-01

    This paper provides an overview of the objectives and status of the Hyper-X program, which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7. Extensive risk reduction activities for the first flight are completed, and non-recurring design activities for the Mach 10 X-43 (third flight) are nearing completion. The Mach 7 flight of the X-43, in the spring of 2001, will be the first flight of an airframe-integrated scramjet-powered vehicle. The Hyper-X program is continuing to plan follow-on activities to focus an orderly continuation of hypersonic technology development through flight research.

  19. Hyper-X Program Status

    NASA Technical Reports Server (NTRS)

    McClinton, Charles R.; Rausch, Vincent L.; Sitz, Joel; Reukauf, Paul

    2001-01-01

    This paper provides an overview of the objectives and status of the Hyper-X program, which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7. Extensive risk reduction activities for the first flight are completed, and non-recurring design activities for the Mach 10 X-43 (3rd flight) are nearing completion. The Mach 7 flight of the X-43, in the spring of 2001, will be the first flight of an airframe-integrated scramjet-powered vehicle. The Hyper-X program is continuing to plan follow-on activities to focus an orderly continuation of hypersonic technology development through flight research.

  20. NASA's Hyper-X Program

    NASA Technical Reports Server (NTRS)

    Rausch, Vincent L.; McClinton, Charles R.; Sitz, Joel; Reukauf, Paul

    2000-01-01

    This paper provides an overview of the objectives and status of the Hyper-X program which is tailored to move hypersonic, airbreathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The first Hyper-X research vehicle (HXRV), designated X-43, is being prepared at the Dryden Flight Research Center for flight at Mach 7 in the near future. In addition, the associated booster and vehicle-to-booster adapter are being prepared for flight and flight test preparations are well underway. Extensive risk reduction activities for the first flight and non-recurring design for the Mach 10 X-43 (3rd flight) are nearing completion. The Mach 7 flight of the X-43 will be the first flight of an airframe-integrated scramjet-powered vehicle.