Showing papers on "Scramjet published in 2004"

A Century of Ramjet Propulsion Technology Evolution

Abstract: A general review is presented of the worldwide evolution of ramjet propulsion since the Wright brothers e rst turned man’ s imagination to e y into a practical reality. A perspective of the technological developments from subsonic to hypersonic e ight speeds is provided to allow an appreciation for the advances made internationally from the early 1900s to current times. Ramjet, scramjet, and mixed-cycle engine types, and their operation and rationale for use are considered. The development history and principal contributing development programs are reviewed. Major airbreathing technologies that had signie cant impact on the maturation of ramjet propulsion and enabled engine designs to mature to their current state are identie ed. The general state of e ight-demonstrated technology is summarized and compared with the technology base of 1980. The current status of ramjet/scramjet technology is identie ed. Ramjet and scramjet propulsion technology has matured dramatically over the years in support of both military and space access applications, yet many opportunities remain to challenge future generations of explorers.

Mixing and combustion studies using cavity-based flameholders in a supersonic flow

Abstract: An experimental investigation of the mixing and combustion processes that occur in and around a cavity-based flameholder in a supersonic flow is reported. Cavity-based flameholders are commonly found in hydrocarbon-fueledscramjet combustors; however, detailed information concerning the behavior of these devices, their optimal shape and fueling strategies, combustion stability, and interactions with disturbances in the main airflow (i.e., shock trains or shock-boundary layer interactions) is largely unavailable in the existing literature. This work is part of an ongoing research program aimed at providing information to help fill these voids and improve the overall understanding of cavities for use as scramjet flameholders.

Mach 4 Performance of Hypersonic Inlet with Rectangular-to-Elliptical Shape Transition

Abstract: Wind-tunnel testing of a hypersonic inlet with rectangular-to-elliptical shape transition has been conducted at Mach 4.0. This fixed geometry inlet had a geometric contraction ratio of 4.8 and was designed using a quasi-streamline tracing technique to have a design point of Mach 5.7. These tests were performed to investigate the starting and backpressure limits of the inlet at conditions well below its design point. Results showed that the inlet required side spillage holes in order to self-start at Mach 4.0. Once started, the inlet generated a compression ratio of 12.6, captured almost 80% of available air and withstood a backpressure ratio of 30.3 relative to tunnel static pressure. The spillage penalty for self-starting was estimated to be 3.4% of available air. These experimental results, along with previous experimental results at Mach 6.2, indicate that fixed-geometry inlets with rectangular-to-elliptical shape transition are a viable configuration for airframe-integrated scramjets that operate over a significant Mach-number range.

Scramjet Inlet Control by Off-Body Energy Addition: A Virtual Cowl

Abstract: The theory of energy addition to hypersonic airflow off the vehicle to increase air mass capture and reduce spillage in scramjet inlets at Mach numbers below the design value is explored. The heated region creates a virtual cowl and deflects flow streamlines into the inlet. Optimization studies are performed with a two-dimensional inviscid fluid code. The best location of the energy addition region is near the intersection of the nose shock of the vehicle with the continuation of the cowl line, and slightly below that line. In that case, the shock generated by the heating is close to the shock that is a reflection of the vehicle nose shock off the imaginary solid surface-extension of the cowl. Effects of the size and shape of the energy addition region on inlet performance are also studied.

Ground Testing of the HyShot Supersonic Combustion Flight Experiment in HEG and Comparison with Flight Data

Abstract: The first phase of the HyShot supersonic combustion ramjet (scramjet) flight experiment program of The University of Queensland in Australia was designed to provide benchmark data on supersonic combustion for a flight Mach number of approximately M=8. The second flight of the HyShot program, performed on July 30th 2002, was successful and supersonic combustion was observed along the specified trajectory range. The operating range of the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Centre (DLR) was recently extended. The facility now has the capability of testing a complete scramjet engine with internal combustion and external aerodynamics at M=7.8 flight con-ditions in altitudes of about 30 km. A post-flight analysis of the HyShot flight experiment was performed using an operational scramjet wind tunnel model with a geometry which is identical to that of the flight configuration.

Analysis of Magnetohydrodynamic Control of Scramjet Inlets

Abstract: Magnetohydrodynamic (MHD) control of forebody flow compression and air mass capture in scramjet inlets is analyzed for flight at Mach 5‐10. Because of the low static temperature, nonequilibrium electrical conductivity is created by electron beams injected into the gas along magnetic field lines. Two-dimensional inviscid steadystate flow equations are solved jointly with equations describing electron-beam-induced ionization profiles, plasma kinetics, and MHD effects. Among several scenarios considered, the scenario with an MHD accelerator has only disadvantages. A modest increase in mass capture can in principle be accomplished with a Faraday MHD generator, if the magnetic field has components both parallel and orthogonal to the flow. The principal focus is on MHD inlet control at flight Mach numbers higher than the design value. The shocks that would otherwise enter the inlet can be moved back to the cowl lip by placing an MHD generator at one of the compression ramps. Analysis shows that the best performance of such a device is achieved with a very short MHD region placed as far upstream (close to the vehicle nose) as possible, in conjunction with a high-current ionizing electron beam. An MHD energy bypass scenario with on-ramp MHD generator for inlet control is briefly discussed.

Contribution to scramjet active cooling analysis using n-dodecane decomposition model as a generic endothermic fuel

Abstract: It is common knowledge that one of the main issues of hypersonic flight is the thermal management of the overall vehicle and more specifically the cooling of the engine, since even composite materials can't withstand the large heat load found in a Scramjet combustion chamber. The other critical point is that mixing and combustion should be sufficiently fast in order to avoid long combustion chamber caused by supersonic internal flow and short residence time. This dual requirement leads to the idea of using inboard fuel as coolant, in order to limit the weight of the cooling system and to benefit from some fuel endothermic capability, i.e. the capability of heavy hydrocarbon fuels to decompose into lighter species, mainly methane, ethane, hydrogen and alkenes from C2 to C10, providing welcomed additional heat sink. Analytical (such as NANCY code) and multidimensional tools (such as CFD-ACE), extensively used for design and analysis of metallic and composite cooled structures are also very efficient to address hydraulic and thermal phenomenon, but, up to now, these approaches did not compute the chemical decomposition of the fuel, except by the way of assumed thermo -dynamical properties. Due to the large number of components found in a traditional jet fuel, the characterization of the thermal decomposition of such products is not a simple issue. In order to obtain a good understanding of the various phenomenon inside a cooled panel of a Scramjet, we selected n-dodecane as a representative surrogate for envisioned endothermic fuel. The thermal decomposition model for n-dodecane was developed by the DCPR and validated thanks to available experimental data. This model, written in a CHEMKIN II format, has been proved able to describe the behavior of n-dodecane between 600 1100 K and 110 bar and could certainly be extrapolated in a wider temperature and pressure range, without taking specifically into account the supercritical phase. This paper will present some numerical results obtained for a cooled composite panel under different heat loads, from a physical point of view (temperature of the panel) and a chemical point of view (composition of the decomposed fuel). INTRODUCTION In a large part of the flight regime, the air-breathing mode appears to be a good solution for future Reusable Space Launchers (RSL). Dual-mode ramjets have been studied to propel such TSTO (Two-Stage To Orbit) or Single -Stage To Orbit (SSTO) vehicles. For example, in the scope of the French PREPHA program, the study of a generic SSTO vehicle led the conclusion that the best type of air-breathing engine could be the Dual-Mode Ramjet (DMR, with subsonic then s upersonic combustion). Airbreathing launchers could typically use hydrogen-fuelled DMR. Less energetic fuels like hydrocarbons could also be used at a Mach number of flight lower than 8, to take advantage of their higher density. In this case, the engine must be able to manage two different fuels. DMR or scramjets designed for hypersonic military applications are typically associated with liquid hydrocarbons and with a maximum flight Mach number of 8 (missiles). Given the very high heat load inside the combustion chamber, active cooling seems to be the only way for a proper thermal management of the engine (Figure 1).

Three-Component Force Measurements on a Large Scramjet in a Shock Tunnel

Abstract: A stress-wave force balance for measurement of thrust, lift, and pitching moment on a large scramjet model (40 kg in mass, 1.165 in in length) in a reflected shock tunnel has been designed, calibrated, and tested. Transient finite element analysis was used to model the performance of the balance. This modeling indicates that good decoupling of signals and low sensitivity of the balance to the distribution of. the load can be achieved with a three-bar balance. The balance was constructed and calibrated by applying a series of point loads to the model. A good comparison between finite element analysis and experimental results was obtained with finite element analysis aiding in the interpretation of some experimental results. Force measurements were made in a shock tunnel both with and without fuel injection, and measurements were compared with predictions using simple models of the scramjet and combustion. Results indicate that the balance is capable of resolving lift, thrust, and pitching moments with and without combustion. However vibrations associated with tunnel operation interfered with the signals indicating the importance of vibration isolation for accurate measurements.

Scramjet Experiments using Radical Farming

Abstract: Scramjet engines are the focus of considerable interest for propulsion in the hypersonic flow regime. One of the serious technical challenges for developing scramjets is reducing the skin friction drag on the engine. The combustion chamber, in particular, is a major contributor to the skin friction drag because of the high density of the flow through that region. This investigation focuses on reducing the combustion chamber skin friction drag by minimising the surface area and size of the combustion chamber and by employing a novel approach to accomplishing combustion. The first design criterion is addressed by using a single internal-combustor scramjet configuration, as opposed to multiple external combustors, and by injecting the fuel on the intake to reduce the mixing length required in the combustor. The second design criterion refers to the use of a new technique called radical farming. This uses the highly two-dimensional nature of the flow through the engine, which is created by deliberately ingesting the leading edge shocks, to achieve combustion at lower mean static pressures and temperatures than generally expected. A simplified approximate theoretical analysis of the radical farming concept is presented. Experiments were conducted in the T4 free-piston shock tunnel on a scramjet model with a single rectangular constant cross-sectional area combustion chamber. Pressure measurements were taken along the centreline of the intake, combustion chamber and thrust surface and across the model width at three locations. Gaseous hydrogen fuel was injected halfway along the intake at a range of equivalence ratios between zero and one. The combustion chamber height was varied from 20mm to 32mm, which varied the contraction ratio of the engine from 4.1 to 2.9. The experiments were conducted at a stagnation enthalpy of either 3MJ/kg or 4MJ/kg. The nominal 3MJ/kg condition corresponds to Mach 7.9 flight at an altitude of 24km. The majority of the 4MJ/kg experiments were conducted at a nominal condition corresponding to Mach 9.1 flight at an altitude of 32km. A small number of 4MJ/kg experiments were conducted at simulated flight altitudes of between 30 and 38km; the flight Mach number for these experiments was approximately 9.0. Thrust was calculated by integrating the centreline pressure distribution over the area of the thrust surface, assuming that the pressure at any axial location was constant across the engine width. These experimental thrust values were compared with theoretical estimates obtained using a one-dimensional analysis and a quasi-two-dimensional analysis. The comparison provided an indication of the level of completion of combustion in the experiments. The difference in thrust produced as a result of combusting fuel was examined by plotting the incremental specific impulse against equivalence ratio. Experimental and theoretical results agreed best at the higher equivalence ratios. Turbulent boundary layer separation correlations were used to provide reasonable estimates for the equivalence ratio at which the flow choked. The drag on the internal flowpath of the scramjet engine was estimated using the quasi-two-dimensional analysis. This drag estimate was combined with the experimental thrust measurements to provide estimates of the net specific impulse. Positive net specific impulse estimates were obtained above a certain minimum equivalence ratio, which depended on the contraction ratio and the test condition. The engine performance was observed to be highly dependent on the two-dimensional shock structure within the engine. Thrust and specific impulse were observed to decrease with increasing simulated flight altitude, as expected. Positive net specific impulse estimates were obtained at equivalence ratios of approximately one for simulated flight altitudes below 35km. Assuming complete combustion and that an equivalence ratio of one can be reached, the configuration considered in the present study can theoretically reach a net specific impulse of approximately 1000s at the 3MJ/kg condition and 500s at the 4MJ/kg condition. These numbers provide a promising testimonial for the use of this configuration, with modifications, as a more efficient alternative to rocket engines.

Hypersonic Boundary/Shear Layer Transition for Blunt to Slender Configurations - A NASA Langley Experimental Perspective

Abstract: : Heating augmentations and temperature increases resulting from boundary layer/shear layer transition during hypersonic flight through the atmosphere of Earth or other planets impose critical constraints on the design of vehicle thermal protection systems and are well documented in the literature. Laminar-to-turbulent transition effects on local surface heat transfer determine thermal protection system material selection, placement, and thickness. In terms of vehicle performance, transition can influence vehicle aerodynamics and scramjet propulsion system performance. The development of numerical tools for the reliable and rapid prediction of boundary layer transition on complex vehicle shapes, however, continues to be hindered by the lack of a practical capability to model the complex physics associated with the transition process. Therefore, until a credible approach to transition prediction is identified that can be implemented in a rapid assessment framework, vehicle designers will continue to rely heavily on empirically derived transition prediction strategies derived from ground-based measurements. With an emphasis on hypersonic boundary layer transition, the focus of the present paper is to highlight recent aerothermodynamic studies in NASA Langley Research Center's conventional hypersonic wind tunnels (as opposed to high enthalpy, impulse facilities) in support of agency access-to-space and planetary entry programs. Configurations of interest include the Shuttle Orbiter and proposed advanced space transportation concepts (reusable and partially reusable crew launch/return vehicles, single and multiple stage-to-orbit rocket /airbreathing propulsion system concepts, hypersonic cruise vehicles, and planetary aerocapture /entry vehicles). Characterization of surface heating on complex shapes and deflected control surfaces and fluid dynamic phenomenon, such as flow separation and wake closure, are addressed.

Analysis of the Magnetohydrodynamic Energy Bypass Engine for High-Speed Airbreathing Propulsion

Abstract: The performance of the MHD energy bypass air-breathing engine for high-speed propulsion is analyzed in this investigation. This engine is a specific type of the general class of inverse cycle engines. In this paper, the general relationship between engine performance (specific impulse and specific thrust) and the overall total pressure ratio through an engine (from inlet plane to exit plane) is first developed and illustrated. Engines with large total pressure decreases, regardless of cause or source, are seen to have exponentially decreasing performance. The ideal inverse cycle engine (of which the MHD engine is a sub-set) is then demonstrated to have a significant total pressure decrease across the engine; this total pressure decrease is cycle-driven, degrades rapidly with energy bypass ratio, and is independent of any irreversibility. The ideal MHD engine (inverse cycle engine with no irreversibility other than that inherent in the MHD work interaction processes) is next examined and is seen to have an additional large total pressure decrease due to MHD-generated irreversibility in the decelerator and the accelerator. This irreversibility mainly occurs in the deceleration process. Both inherent total pressure losses (inverse cycle and MHD irreversibility) result in a significant narrowing of the performance capability of the MHD bypass engine. The fundamental characteristics of MHD flow acceleration and flow deceleration from the standpoint of irreversibility and second-law constraints are next examined in order to clarify issues regarding flow losses and parameter selection in the MM modules. Severe constraints are seen to exist in the decelerator in terms of allowable deceleration Mach numbers and volumetric (length) required for meaningful energy bypass (work interaction). Considerable difficulties are also encountered and discussed due to thermal/work choking phenomena associated with the deceleration process. Lastly, full engine simulations utilizing inlet 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.

Conceptual study of a combined cycle engine for an aerospace plane

Abstract: Operating conditions and performances of a fixed-geometry combined-cycle engine for a single-stage-to-orbit aerospace plane were calculated with a simple simulation model. With the flow conditions calculated with the model, the cooling requirement of the engine and pitching moment of the plane were investigated. The engine was composed of an ejector-jet mode, a ramjet mode, a scramjet mode, and a rocket mode. The engine had a fixed geometry in its operation. Subsonic combustion was conducted with no second throat in the combustor under the ejector-jet mode and the ramjet mode. Propellants were liquid hydrogen and liquid oxygen. The coolant flow rate became larger than the fuel flow rate. The excessive flow rate decreased the specific impulse above Mach 9 and restricted application of the airbreathing engine mode up to Mach 11. The pitching moment of the plane would be balanced even in the space in the configuration with the combined-cycle engine mounted on the windward surface.

Investigation of an intake injected hot-wall scramjet

Abstract: Injecting fuel on the intake of a scramjet allows the fuel and air to mix before entering the combustion chamber. This allows for a reduction in combustion chamber length, a major contributor to scramjet drag. This investigation aims to check if any combustion is occurring on the intake when the intake wall is hot. Gaseous Hydrogen fuel was injected at an angle of 45 degrees to the freestream approximately halfway along the intake of a two-dimensional scramjet model. The T4 shock tunnel, an impulse testing facility, was used to produce a hypersonic flow with freestream enthalpy of 3.0MJ/kg and Mach number of 6.5. The model was modified such that the intake wall could be heated to temperatures experienced in continuous flight. A pressure rise on the intake or an increase in the initial shock angle after the fuel was injected would indicate the occurrence of intake combustion. Neither of these things was observed in current experiments consistent with no ignition on the intake. The experimental procedure and results are discussed in this paper and compared with theoretical results.

MHD Power Generation in Scramjet Engines in Conjunction with Inlet Control

Abstract: Megawatt-scale MHD power generation on board of air-breathing hypersonic vehicles is made possible by the high flow enthalpy and velocity, and by the high static temperature (2,300 - 3,500 K) in the combustor and just downstream of it. In this paper, a 1D model of MHD power generation downstream of scramjet combustor is developed, and power generation performance is predicted for a variety of generic hydrogen- fueled engine regimes. A modest amount of potassium seed is shown to yield good performance at magnetic fields on the order of 1 Tesla. One possible use of the MHD-generated power would be to add energy to the flow upstream of the inlet in order to optimize the inlet performance. In particular, MHD generated power can be used to create a virtual cowl that would increase mass capture and thrust of the engine. The paper also demonstrates that the megawatt-scale power, when added to the flow upstream of the inlet throat, can reduce the Mach number, leading to the possibility of substantially shortening or even eliminating the isolator.

Comparison of hthl and vthl air-breathing and rocket systems for access to space

Abstract: A design study was performed to define and compare the parameters of horizontal- and verticaltakeoff reusable launch vehicle systems in an effort to identify optimal configurations for improved access to space. The payload requirement for each vehicle was 20,000 pounds delivered to low Earth orbit. The investigation considered both rocket powered two-stage and air-breathing ramjet/scramjet powered single-stage vehicle configurations. All vehicles were first analyzed utilizing liquid hydrogen for the entire trajectory and then re-analyzed with liquid hydrocarbon fuel in the first trajectory segment. The vertical air-breathing vehicles were found to have the lowest empty weights and gross takeoff weights of all the vehicle configurations, with inward turning inlets outperforming the two-dimensional inlets. For the rockets, the lightest empty weight was achieved with the use of hydrocarbon fuel in the booster and hydrogen fuel in the orbiter. The best horizontal takeoff vehicle is the all hydrogen inward-turning airbreather.

Understanding isolator performance operating in the separation-shock mode

Abstract: W = width of isolator duct, in. δ = boundary layer thickness, in. Isolators, also called supersonic diffusers, are a required component of airbreathing scramjet engines and are used to isolate the precombustion pressure rise generated by the combustion process from the supersonic inlet compression. At high inflow Mach number or for isolators with high aspect ratio, the separation shock mode of operation has been observed, which is characterized by strong three-dimensional effects and long duct length requirements to contain a specified pressure rise. In the present paper, the effects of inflow Mach number and aspect ratio on isolator operating characteristics are calculated for rectangular isolators. The results show that the dominant flow feature is a corner flow separation that exists at all inflow Mach numbers and aspects ratios considered. The 3D flow variations are seen to become stronger with both increasing inflow Mach number and aspect ratio.

Piv measurement of a 3-dimensional reacting flow in a scramjet combustion chamber

Abstract: This paper describes the PIV measurements at the M11-4 test facility at DLR-Lampoldshausen. The motivation for these experiments is to gain deeper insight into the flow fields inside a model scramjet combustion chamber, in which one of two hydrogen injectors of different geometry can be installed at a time. These flow fields are expected to vary depending on the injector in use. Other applied diagnostic methods like wall static pressure measurements or OH-LIF reveal differences in the flow in that respect. Successful PIV measurements could deliver the data required to enable a better understanding of how each injector geometry influences the flow, resulting in those differences observed. Problems faced are high flow velocities, strong gradients due to shock waves and the 3-dimensionality of the flow itself. To a large extent, those problems could be overcome and new insights with about the influence of the injector geometry on the flow field were gained by means of the PIV measurements.

Reusable thrust-powered sled mounted on an inclined track for launching spacecraft and airborne vehicles at supersonic speeds

Abstract: This invention will allow a reusable thrust-powered sled mounted on an inclined track to launch spacecraft or airborne vehicles from earth at supersonic speeds using existing technology properly integrated into an inclined track system. If launched up a tunnel track, a rear blast shield can trap the rocket exhaust to provide a pneumatic boost upon launch. The sled can also launch ramjet or scramjet powered vehicles from earth by achieving the Mach 2+ speed necessary to ignite their engines. This system is much safer than the traditional method of launching rockets since weather is less a factor and the launch can be aborted if problems develop. Moreover, it is far less costly since the engines on the sled can be reused hours after a launch and the track can accommodate a variety of sleds to launch objects of many different sizes.

Analysis and Design of Rectangular-Cross-Section Nozzles for Scramjet Engine Testing

Abstract: The flow in the square-cross-section Mach-6 nozzle used in the NASA Langley Research Center Arc-Heated Scramjet Test Facility has been analyzed using three-dimensional viscous CFD methods. The primary cause of the non-uniform flow exiting the nozzle is identified as cross-flow pressure gradients imposed on wall boundary layers. The cross-flow pressure gradients cause the boundary layer to roll up into counter-rotating vortex pairs on each of the four sides of the nozzle. These four vortex pairs produce significant non-uniformity in the nozzle-exit flow. In order to improve the quality of the test flow in the facility, two alternative nozzle designs (one axisymmetric and one rectangular with a 2-D contour) have been investigated. While the axisymmetric design produced the most uniform flow, the 2-D design also produced very good flow. The 2-D design was selected for further refinement, resulting in a new nozzle design which has been constructed and awaits calibration.

Cfd simulations of a 3-d scramjet flameholder using reduced chemical kinetic mechanisms

Abstract: Reduced chemical kinetic mechanisms for combustion of ethylene and n-heptane have been implemented into the VULCAN CFD code and used for simulations of a 3-D scramjet flameholder. A reduced mechanism for JP-8 has been created, implemented into VULCAN and used to model a 2-D supersonic flame. The reduced mechanisms have been created using the CARM (Computer Aided Reduction Method) software. These reduced mechanisms were tested against measurements and detailed chemistry for their ability to predict ignition delay over a range of conditions and found to give good agreement. 3-D high speed CFD simulations using a reduced mechanism for hydrocarbon/air combustion demonstrate the ability of reduced mechanisms to implement realistic hydrocarbon combustion chemistry models into CFD.

Scramjet Optimization for Maximum Off-Design Performance

Abstract: Aspects relating to the design, analysis, and optimization of scramjet engines for peak on and off-design performance are explored in this paper. Tip-to-tail flowpaths are developed and compared using both fixed and variable geometry combustors to assess their off-design penalties for Mach 8 and Mach 10 design points. The performance characterization is done for changes in Mach number, dynamic pressure, angle of attack, and fuel equivalence ratio. A quasi-one-dimensional engine model, including the effects of fuel injection location, mixing length, fuel mixing efficiency, chemical production rates, heat transfer, and viscous losses is utilized to assess the effects of hydrogen finite-rate chemistry. Due to the coupled nature of hypersonic propulsion/airframe integration, realistic two-dimensional inlet and nozzle designs have been included in this analysis.

Predicted scramjet testing capabilities of the proposed RHYFL-X expansion tube

Abstract: For the Supersonic Combustion Ramjet (Scramjet) to be realized as a viable hypersonicnpropulsion option, significant testing must be performed in test flows which duplicate thosenthat would be experienced during anticipated atmospheric flight. The expense associatednwith real flight tests necessitates that ground-based testing play a significant role in thenexperimental development of the scramjet. Based on current, or even near-term technologies,nthe expansion tube concept offers superior physical simulation of hypersonic flow. This studyninvestigates the potential performance capabilities of a large-scale, high-performance free-pistonndriven expansion tube based on the RHYFL shock tunnel project. The predictions arenfocused specifically on the ability of the RHYFL-X expansion tube to generate sub-orbitalnscramjet testing conditions and are obtained through one-dimensional and axisymmetricnsimulations of this proposed facility.n n To validate the simulation techniques and approximations used to model the flow in thenRHYFL-X expansion tube, simulations of a currently operating expansion tube, X2, are presented.nThese one-dimensional and axisymmetric simulations initially assumed equilibriumnchemistry. The driver gas conditions after the two-stage compression process were obtainednvia a combination of numerical and experimental analyses. An accurate knowledge of thendriver length at the point of primary diaphragm rupture was required for sound agreementnwith experimental results. The inertial effect of the secondary diaphragm was also examinednfor different X2 operating conditions. Results show that a hold-time imposed on thensecondary diaphragm improved agreement with experimental data.n n The effects of finite-rate chemistry on the air test gas of two standard operating conditionsnin the X2 facility were also investigated. A 5 species, 17 reaction model of air was used inninviscid one-dimensional simulations of a 6.8km/s condition and a 9.7km/s condition. Byncomparison with equilibrium simulations of the same conditions it was seen that finite-ratenchemistry effects can be significant in conditions where test gas dissociation occurs prior tonthe unsteady expansion. For the higher enthalpy condition, the significant dissociation ofnthe air test gas caused by the faster primary shock, combined with the more severe unsteadynexpansion process, resulted in a test flow highly influenced by nonequilibrium phenomena.nThe low primary shock speeds required to produce atmospheric static temperatures in thenfinal test flow of the proposed RHYFL-X expansion tube was shown to result in negligible, ifnany, dissociation of the test gas prior to expansion. Even assuming the worst case scenario ofna completely frozen expansion, these minimal dissociation levels prior to expansion indicatenthat the three RHYFL-X operating conditions investigated would have test flows essentiallyndissociation free.n n Results from one-dimensional and axisymmetric simulations of the RHYFL-X expansionntube indicate that this proposed facility would be capable of generating true Mach numberntesting conditions over the intended scramjet flight trajectory, while inheriting the short testntimes and limited core flow diameters associated with expansion tubes. Simulated pressuresnwell in excess of those that would be experienced during flight indicate that this facilitynwould offer the unique capability of duplicating freestream conditions required for accuratenaerodynamic, heating and combustion testing of integrated scramjet models. The use of annozzle was also investigated for increasing the diameter of the core test flow. While seeingnmoderate increases in core flow diameter, the viscous axisymmetric simulations also displayednthe desirable characteristic of increasing the duration of steady flow.n

Effects of hydrogen-air non-equilibrium chemistry on the performance of a model scramjet thrust nozzle

Abstract: Two aspects of hydrogen-air non-equilibrium chemistry related to scramjets are nozzle freezing and a process called 'kinetic afterburning' which involves continuation of combustion after expansion in the nozzle. These effects were investigated numerically and experimentally with a model scramjet combustion chamber and thrust nozzle combination. The overall model length was 0.5m, while precombustion Mach numbers of 3.1 +/- 0.3 and precombustion temperatures ranging from 740K to 1,400K were involved. Nozzle freezing was investigated at precombustion pressures of 190kPa and higher, and it was found that the nozzle thrusts were within 6% of values obtained from finite rate numerical calculations, which were within 7% of equilibrium calculations. When precombustion pressures of 70kPa or less were used, kinetic afterburning was found to be partly responsible for thrust production, in both the numerical calculations and the experiments. Kinetic afterburning offers a means of extending the operating Mach number range of a fixed geometry scramjet.

CFD Capabilities for Hypersonic Scramjet Propulsive Flowpath Design

Abstract: This paper discusses the status of our CFD capabilities for scramjet flow path analysis and design, as well as key enabling technologies that enhance accuracy and efficiency. A scramjet engine flow path encompasses some of the most complex aero-propulsive physics including; laminar to turbulent flow transition, complex shock interactions, and multi-species reacting chemistry. Through the judicious use of advanced modeling and adaptive procedures to ensure adequate resolution, CFD can now predict the performance characteristics of proposed engine designs. Several enhancements to the baseline modeling approach are presented and their influence is demonstrated on unit problems representative of scramjet flow paths. Nomenclature Pr t = Turbulent Prandtl Number Sc t = Turbulent Schmidt Number Le = Lewis Number, Pr t /Sc t Re θ = Reynolds Number based on momentum boundary layer thickness M = Mach Number k,ε = Turbulent kinetic energy, turbulent dissipation rate T w = Temperature of wall PDE = Partial differential equation

Modeling of Plasma Virtual Shape Control of Ram/Scramjet Inlet and Isolator

Abstract: The paper explores a concept of ram/scramjet propulsion control by energy addition and extraction in the propulsion flowpath and the reverse energy bypass concept. Instead of variable geometry, the concept relies on virtual shapes created by plasma/MHD devices or by other methods (including plasma-controlled external combustion). An inherent advantage of the proposed plasma/MHD control system is its flexibility, fast response, and the absence of moving parts. The fixed geometry is optimized for Mach 7 flight. At Mach numbers higher than the design value, an MHD generator placed at the first compression ramp and using ionization by electron beams can restore the shock-on-lip condition, while operating in self-powered regime. The magnetic field of 1.5-1.7 Tesla would be sufficient for Mach 8 flight. At Mach numbers below the design value, inlet performance can be controlled by energy addition, with the power supplied by an MHD generator placed downstream of the combustor. This concept is called the reverse energy bypass. In one scenario, the inlet flow spillage can be reduced by Virtual Cowl – a heated region placed upstream of the cowl and slightly below it. With optimally located Virtual Cowl, calculations with conservative assumption regarding power transmission losses show that the reverse bypass can increase thrust by about 10% at Mach 6. In another scenario, distributed heating of the flow upstream of the inlet throat in the ramjet regime (Mach 4-6), with the heating rate of about 6.3-8.5% of the total enthalpy flux, can bring the throat Mach number close to 1, thus making the isolator duct virtually unnecessary. Although the reverse bypass system with inlet heating would reduce thrust by about 16% at Mach 5, the performance penalty at the vehicle acceleration stage can be offset by the increased efficiency during the cruise due to the absence of weight and cooling burden normally caused by the long isolator duct.