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NASA NTRS2002

Dual-Mode Combustion

The Department of Mechanical and Aerospace Engineering at the University of Virginia has conducted an investigation of the mixing and combustion processes in a hydrogen fueled dual-mode scramjet combustor. The experiment essentially consisted of the "direct connect" continuous operation of a Mach 2 rectangular combustor with a single unswept ramp fuel injector. The stagnation enthalpy of the test flow simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and laser based diagnostics. These diagnostics included, pressure and wall temperature measurements, Fuel Plume Imaging (FPI) and Particle Image Velocimetry (PIV). A schematic of the combustor configuration and a summary of the measurements obtained are presented. The experimental work at UVa was parallel by Computational Fluid Dynamics (CFD) work at NASA Langley. The numerical and experiment results are compared in this document.

Goyne, Christopher P.↗

NASA NTRS2012

Computational Analyses of the LIMX TBCC Inlet High-Speed Flowpath

Reynolds-Averaged Navier-Stokes (RANS) simulations were performed for the high-speed flowpath and isolator of a dual-flowpath Turbine-Based Combined-Cycle (TBCC) inlet using the Wind-US code. The RANS simulations were performed in preparation for the Large-scale Inlet for Mode Transition (LIMX) model tests in the NASA Glenn Research Center (GRC) 10- by 10-ft Supersonic Wind Tunnel. The LIMX inlet has a low-speed flowpath that is coupled to a turbine engine and a high-speed flowpath designed to be coupled to a Dual-Mode Scramjet (DMSJ) combustor. These RANS simulations were conducted at a simulated freestream Mach number of 4.0, which is the nominal Mach number for the planned wind tunnel testing with the LIMX model. For the simulation results presented in this paper, the back pressure, cowl angles, and freestream Mach number were each varied to assess the performance and robustness of the high-speed inlet and isolator. Under simulated wind tunnel conditions at maximum inlet mass flow rates, the high-speed flowpath pressure rise was found to be greater than a factor of four. Furthermore, at a simulated freestream Mach number of 4.0, the high-speed flowpath and isolator showed stability for freestream Mach number that drops 0.1 Mach below the design point. The RANS simulations indicate the yet-untested highspeed inlet and isolator flowpath should operate as designed. The RANS simulation results also provided important insight to researchers as they developed test plans for the LIMX experiment in GRC s 10- by 10-ft Supersonic Wind Tunnel.

Dippold, Vance F., III↗

NASA NTRS2012

Dual-Mode Scramjet Flameholding Operability Measurements

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 sensitivity of flameholding limits to inlet air vitiation by using electrically heated air and adding steam at levels to simulate vitiated air heaters. The measurements showed no significant difference in blowout temperature with inlet air mole fractions of steam from 0 to 6.7%.

airbreathing engines aircraft↗

NASA NTRS1973

Emission calculations for a scramjet powered hypersonic transport

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.

Lezberg, E. A.↗

NASA NTRS2015

Development of a Premixed Combustion Capability for Scramjet Combustion Experiments

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.

Rockwell, Robert D.↗

NASA NTRS2008

Forebody and Inlet Design for the HIFiRE 2 Flight Test

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

Ferlemann, Paul G.↗

NASA NTRS1992

NASA's hypersonic propulsion program: History and direction

Research into hypersonic propulsion; i.e., supersonic combustion, was seriously initiated at the Langley Research Center in the 1960's with the Hypersonic Research Engine (HRE) project. This project was designed to demonstrate supersonic combustion within the context of an engine module consisting of an inlet, combustor, and nozzle. In addition, the HRE utilized both subsonic and supersonic combustion (dual-mode) to demonstrate smooth operation over a Mach 4 to 7 speed range. The propulsion program thus concentrated on fundamental supersonic combustion studies and free jet propulsion tests for the three dimensional fixed geometry engine design to demonstrate inlet and combustor integration and installed performance potential. The developmental history of the program is presented. Additionally, the HRE program's effect on the current state of hypersonic propulsion is discussed.

Wander, Steve↗

NASA NTRS2015

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

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.

Cantu, Luca M. L.↗

NASA NTRS1978

Conceptual study of hypersonic airbreathing missiles

The purpose of this paper is to report recent results of an in-house conceptual study to evaluate the performance potential and research needs of airbreathing hypersonic missiles. An alkylated-borane (noncryogenic) fueled, dual-mode, ramjet/scramjet propulsion system structured with a Rene 41 inlet and a carbon-carbon combustor was assumed along with a Lockalloy heat sink fuselage structure and beryllium wings and control surfaces. Performance for an air-launched baseline missile with a 961 pound staging weight containing a 100 pound payload indicated excellent long range cruise, moderate acceleration and high maneuverability potential. A sizing study indicates that Mach 6 cruise ranges of the order of 2500 nautical miles for payloads of 300 pounds can be achieved with moderate size missile carry weights (9000 lbs.). Aerodynamic heating analyses indicate that unprotected heat-sink structures with internal insulation are feasible for ranges of several hundred miles. For ranges of several thousands of miles a multiwall radiation shield (Inconel/titanium) was selected for protection of the internally insulated heat sink structure.

Hunt, J. L.↗

NASA NTRS2012

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

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.

Cutler, Andrew D.↗

NASA NTRS2012

OH PLIF Visualization of the UVa Supersonic Combustion Experiment: Configuration A

Hydroxyl radical (OH) planar laser-induced fluorescence (PLIF) measurements were performed in the University of Virginia s dual-mode scramjet experiment. The test section was set up in configuration A, which includes a Mach 2 nozzle, combustor, and extender section. Hydrogen fuel was injected through an unswept compression ramp at two different equivalence ratios. Through the translation of the optical system and the use of two separate camera views, the entire optical range of the combustor was accessed. Single-shot, average, and standard deviation images of the OH PLIF signal are presented at several streamwise locations. The results show the development of a highly turbulent flame structure and provide an experimental database to be used for numerical model assessment.

Johansen, Craig T.↗

NASA NTRS1987

Ground tests confirm the promise of hypersonic propulsion

The feasibility of developing a hypersonic hydrogen-fueled airbreathing vehicle capable of Mach 8 is examined. A modular dual-mode scramjet concept integrated with the airframe has been proposed. The research involved in the formation of this design is discussed. Particular attention is given to studies concerned with airframe integration; the evaluation of hypersonic inlets; and the analysis of viscous phenomena. The injector and combustor configurations of the modular engine, which is designed to operate as a ramjet and scramjet, are analyzed. The engine thrust performance of subscale engine modules was tested, and positive results were obtained. Simulation testing of a full-size model is being planned.

Anderson, Griffin Y.↗

NASA NTRS2002

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

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.

Cockrell, Charles E., Jr.↗

NASA NTRS1997

Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet

Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

Source record↗

NASA NTRS2002

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

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.

Waltrup, Paul J.↗

NASA NTRS1999

CIAM/NASA Mach 6.5 Scramjet Flight and Ground Test

The Russian Central Institute of Aviation Motors (CIAM) performed a flight test of a CIAM-designed, hydrogen-cooled/fueled dual-mode scramjet engine over a Mach number range of approximately 3.5 to 6.4 on February 12, 1998, at the Sary Shagan test range in Kazakhstan. This rocket-boosted, captive-carry test of the axisymmetric engine reached the highest Mach number of any scramjet engine flight test to date. The flight test and the accompanying ground test program, conducted in a CIAM test facility near Moscow, were performed under a NASA contract administered by the Dryden Flight Research Center with technical assistance from the Langley Research Center. Analysis of the flight and ground data by both CIAM and NASA resulted in the following preliminary conclusions. An unexpected control sensor reading caused non-optimal fueling of the engine, and flowpath modifications added to the engine inlet during manufacture caused markedly reduced inlet performance. Both of these factors appear to have contributed to the dual-mode scramjet engine operating primarily in a subsonic combustion mode. At the maximum Mach number test point, combustion caused transition from supersonic flow at the fuel injector station to primarily subsonic flow in the combustor. Ground test data were obtained at similar conditions to the flight test, allowing for a meaningful comparison between the ground and flight data. The results of this comparison indicate that the differences in engine performance are small.

Voland, R. T.↗

NASA NTRS2010

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

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 identified.

Hass, Neal E.↗

NASA NTRS2011

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

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.

Cabell, Karen↗