Showing papers on "Scramjet published in 2000"

Aerodynamic Database Development for the Hyper-X Airframe Integrated Scramjet Propulsion Experiments

Abstract: This paper provides an overview of the activities associated with the aerodynamic database which is being developed in support of NASA''s Hyper-X scramjet flight experiments. Three flight tests are planned as part of the Hyper-X Program. Each will utilize a small, non-recoverable research vehicle with an airframe integrated scramjet propulsion engine. The research vehicles will be individually rocket boosted to the scramjet engine test points at Mach 7 and Mach 10. The research vehicles will then separate from the first stage booster vehicle and the scramjet engine test will be conducted prior to the terminal decent phase of the flight. An overview is provided of the activities associated with the development of the Hyper-X aerodynamic database, including wind tunnel test activities and parallel CFD analysis efforts for all phases of the Hyper-X flight tests. A brief summary of the Hyper-X research vehicle 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. Brief comments on the planned post flight data analysis efforts are also included.

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

Abstract: 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 cowl-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.

Comparison of Supersonic Combustion Between Impulse and Vitiation-Heated Facilities

Abstract: A comparison has been made between supersonic combustion in two commonly used, but fundamentally different, facilities for scramjet research, a vitiation-heated blowdown tunnel and a free-piston shock tunnel. By passing the shock-tunnel freestream flow through a normal shock and then expanding it to Mach 2.5, combustor inlet conditions and geometries were nominally replicated between the two facilities. A constant-area rectangular duct and a diverging duct, both employing central-strut hydrogen injection, were used. Boundary-layer separation and choking in the constant-area duct limited supersonic combustion comparisons up to a fuel equivalence ratio of the order of 0.3. The experimental results also show that the onset of boundary-layer separation occurs at the same combustor pressure loads and that it behaves similarly in the different facilities. With the diverging duct, comparisons were made up to an equivalence ratio of 1.05. Agreement between the results obtained in the two facilities is within experimental error when the different freestream and boundary layers are accounted for.

Oh planar laser-induced fluoroescence imaging in a hydrocarbon-fueled scramjet combustor

Abstract: Planar laser-induced fluorescence (PLIF) imaging of OH has been completed tthrough the piloting section of a hydrocarbon-fueled scramjet. This pilot consists of flush-wall fuel injection followed by a recess, or cavity, in one wall. Images were obtained for both gaseous (ethylen) and liquid (JP-7) fuel combustion. For the gaseous-fuel tests, ethylene was introduced through four flush-wall, low-angle injectors placed upstream of the cavity. For the liquid-fuel tests, injection was normal to the crossflow through seven injectors (four in the bottom, wall, three in the top wall). Introducing a small amount of gas into the liquid in the bottom wall injectors enhanced atomization of the liquid column. Flight conditions between March 4 and 5 and dynamic pressures between 23.9 and 71.7 kPa are simulated. Instantaneous images show the dynamis of the combustion process, suggest the process is premixed in nature, and reveal the presence of large-scale structures. Average images at different axial locations show the effects of total temperature and dynamic pressure on the combustion process. Increasing temperature broadens the time-averaged flame zone, while increasing dynamic pressure tends to force the flame against the combustor sidewall. At a given axial location, the time-averaged reaction zone for ethylene is larger than that for JP-7.

Fuel Injection into a Supersonic Airflow by Means of Pylons

Abstract: The mixing and reaction processes in a scramjet combustion chamber have been experimentally investigated. Hydrogen was injected into a preheatedMach 2.15 airstream by means of pylon-like fuel injectors. The supersonic  ame was stabilized in a purely reaction–kinetical way; i.e., by means of self-ignition. Various pylon designs have been employed to study their in uence on fuel mixing and combustion in the supersonic airstream. Based on the results, an optimized fuel injector has been designed and tested. To assess the reacting  ow, nonintrusive, optical measurement techniques have been employed: the Rayleigh scattering technique to study the injected mixing jets, and the self- uorescence of the OH radical to determine location and intensity of the reaction zones. Additionally, the wall static pressure has been measured.

Self-ignition and transition to flame-holding in a rectangular scramjet combustor with a backward step

Abstract: Self-ignition and transition to flame-holding in a rectangular scramjet combustor with a backward step have been investigated experimentally and numerically in order to clarify whether these are dominated by the near-field phenomena or the far-fiel phenomena Hydrogen fuel was injected perpendicularly into the Mach 20 high-enthalpy airflow downstram of the step The details of the flowfield were captured by a three-dimensional full Navier-Stokes numerical code with a large-eddy simulation turbulence model and a detailed chemical reaction model The characteristics of self-ignition and transition to a bulk flame are explained by the Damkohler number (Da) Da in the combustor is lower than unity right after the fuel injection due to the low pressure level, and the flowfield in the combustor is essentially reaction-limited therefore, fast chemical reaction is allowed only in the hot boundary layer and the recirculation zone behind the step Although self-ignition was observed either in the near-field or in the far-field boundary layers according to the condition of the mixture, intensive chemical reaction was not observed outside the boundary layer at the early step because of the very low Da and therefore the transition to the bulk flame was difficult However, the transition to the bulk flame was achieved by the propagation of the shock wave system which increased the Da by 1 or 2 orders of magnitude and turns the flowfield to mixing-limited The shock wave was triggered by the slight increase of the pressure level in the far field The generation and the propagation of the shock wave were significantly affected by the combustor geometry and the heat release in the far field: thus, the transition was essentially dominated by the far-field phenomena

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

Abstract: Airframe-integrated scramjet engine tests have been completed at Mach 7 in 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.

Self-Ignition and Supersonic Reaction of Pylon-Injected Hydrogen Fuel

Abstract: The e ame stabilization and reaction processes in a scramjet combustion chamber have been experimentally investigated. Hydrogen was injected into a vitiated Mach 2.15 airstream by means of pylon-like fuel injectors. The supersonice amewasstabilizedina purelykineticalway;i.e.,bymeansoffuelself-ignition.Thee ame-stabilization mechanisms have been studied. The interaction between the gasdynamics and the reaction kinetics are discussed. A small wedgehas beenmounted into thetestcombustorto modify theobliqueshockstructure.Theresponseofthe reacting e ow on this changehas been observed. To assessthereacting e ow, only nonintrusive, optical measurement techniques have been employed: the schlieren technique to visualize the e ow structure, the Rayleigh scattering technique to study the injected mixing jets, as well as the self-e uorescence of the OH radical to determine location and intensity of the reaction zones. Additionally, the wall static pressure has been measured.

RBCC Engine-Airframe Integration on an Osculating Cone Waverider Vehicle

Abstract: An aerodynamic model for an RBCC engine-airframe integrated vehicle using an osculating cone waverider forebody is presented. Emphasis is placed on the hypersonic portion of the vehicle trajectory, where scramjet operation occurs. An engine model that includes finite-rate chemistry of hydrogen and air is developed that predicts the flowfield throughout the combustor. Issues of off-design Mach number, angle of attack, and performance through a trajectory are addressed for a non-optimal example vehicle.

Numerical Investigation of Hypersonic Intake Flows

Abstract: A numerical and experimental analysis of scramjet intake flows has been initiated at RWTH Aachen University as part of the Research Training Group GRK 1095: “Aero-Thermodynamic Design of a Scramjet Engine for a Future Space Transportation System”. This report presents an overview of the ongoing work on the numerical simulations of air intake flow using two different, well validated Reynolds averaged Navier Stokes solvers. Several geometry concepts e.g. 2D intake, 3D intake using a single or double ramp configuration were investigated. One example for the so-called 2D intake can be seen in Fig. 1 and for a 3D intake in Fig. 2. To analyze the effects these different geometries have on the flow, especially on the separation bubble in the isolator inlet as well as on transition and efficiency, several numerical simulations (2D and 3D) were performed using a variety of turbulence models. Mostly the Spalart–Allmaras – one equation model and the so called SSG–Reynolds stress model by Speziale, Sakar and Gatski were used. The data obtained will be compared with experimental results. These experiments started in March 2007. It has to be said that not all results presented here were achieved using the NEC computing cluster. For comparison several calculations were conducted on the IBM Jump system of the Julich Research Centre and on the SUN cluster of RWTH Aachen University. At the end of this report we give comments on the computational performance.

MHD Energy Bypass Scramjet Performance with Real Gas Effects

Abstract: 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.

A Streamlined. Pressure-Matched Fuel Injector for Scramjet Application

Abstract: Pre-test and post-test computational fluid dynamics (CFD) data have been generated for an experimental scramjet fuel injector concept. These cold-flow calculations are part of an experimental and computational evaluation of an innovative, flush-mounted fuel injector designed by Aerojet General Corporation to provide maximum penetration with minimal losses. The key to the success of the design is the concept of local pressure matching. Results show fuel penetration far in excess of a priori expectations for the test conditions but reasonable agreement between CFD and experiment. As a result of excessive penetration, shock strength is higher than expected and total pressure losses are also high. Mixing efficiencies are comparable to those produced by an equivalent normal circular injector. Data obtained from CFD have proven beneficial to the experimental program by alerting researchers to potential problems, illuminating flow features inaccessible by current testing methods, and guiding the selection of fuel pressures for the achievement of local pressure matching.

Real-Gas Calculation of MHD-Bypass Scramjet Performance

Abstract: The theoretical performance of a scramjet propulsion system incorporating an magneto-hydrodynamic (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 finiterate 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 nonMHD 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.

Criteria for flame holding in H2-fueled scramjet engines

Abstract: Although self-sustaining combustion has often been termed auto ignition in studies on scramjet combustors, it should be interpreted as a flame-holding phenomenon rather than ignition. Criteria for flame holding in H2-fueled scramjet engines have been discussed by approximating the recirculation zone in the combustor with a perfectly stirred reactor. The influence of mass flow rates and reactor volumes on flame holding can be summarized by a critical reaction time in the recirculation zone. It was found that the reaction time was proportional to pressure, with P−13 for H2 reactions, and the recirculation zone should be occupied with reactants having equivalence ratios in the range 0.4

A study of a new ramp injector for improvement of supersonic mixing

Abstract: A study on the mixing of air and the secondary gas is conducted experimentally in order to develop the design of SCRAM-jet engine. Two methods of injection are compared. In the present study, sonic Helium gas is injected into supersonic air flow of A4, = 4.0. Flow visualization (Schlieren photographs and oil flow technique) and volume fraction measurements are conducted in the experiments. From the volume fraction measurements, it is turned out that the injection from the upper face of the ramp injector is superior in the supersonic mixing to the injection from the rear face of the ramp injector.

Current status and potentialities of wind tunnels with MHD acceleration

Abstract: The activities aimed at development of hypersonic wind tunnels with MHD gas accelerator started in the late 1950s, when it became clear that the uses of hypersonic wind tunnels of the classical type employing the principle of adiabatic expansion for gas acceleration are subject to limitations. A flow train was developed for such a wind tunnel including a source of conducting gas (electric-arc heater, region behind the reflected shock wave, high-pressure chamber of an impulse wind tunnel, etc.), a system for delivery of readily ionizing seed (K, Na, Cs, and their compounds), a primary supersonic nozzle with a Mach number M ≃1.6-3.0, the MHD-accelerator channel proper, the working section, and a gas exhaust system. In the United States, these studies were performed at the NASA Langley Center, AEDC, and General Electric; in Russia, at the Central Institute of Aerohydrodynamics (TsAGI). Some work was done in France at ONERA and in Germany. This review contains an analysis of various design solutions of the MHD channel proper, seed delivery systems, and systems of electric supply and magnetic field generation, as well as a comparison of the procedures for testing and acquisition and treatment of experimental data. Special note is made of the difficulties confronting the researchers and of the reasons for simultaneous cessation of research activities in the United States at almost all research centers. It is indicated that in Russia, unlike the United States, much more extensive studies were made into the physics of combustion of discharge in gas flows at Mach numbers M = 0–4.5 both with and without a magnetic field. Data are given on the behavior of current distribution in the electrodes, propagation of microarc discharges, boundary layer separation, and magnitude of heat fluxes with the flow parameters corresponding to those of the MHD accelerator. The basic characteristics of the TsAGI hypersonic wind tunnel are given along with the results of its utilization for solving practical problems of aerodynamics. Results are also given of the investigation of a high-enthalpy MHD generator in which the flow from the MHD accelerator is used as the working medium. The advantages are demonstrated of using facilities with MHD gas accelerator in solving the problems of development and testing of scramjet engines (SJE) for transatmospheric flying vehicles, and information is given about the possible parameters (gasdynamic and electrodynamic) of such facilities and their designs. The main problems are listed that must be solved in developing hypersonic facilities with MHD gas accelerator.

Mixing and Combustion in Scramjet Combustor with Raised and Relieved Ramp

Abstract: A numerical study of wall-mounted parallel injector ramp has been conducted to investigate the mixing process of scramjet engine using unstructured grids. Two different ramp configurations have been studied, raised and relieved ramps. A three dimensional models for the two different rampconfigurations have been used with two different side sweep angles, 0 (unswept) and 5 degrees. Numerical results are obtained using an existing CFD code "FLUENT" with unstructured grids with a size approximately 300,000 nodes. Results are presented for both non-reacting and reacting flows. Results show that the unswept relieved ramp gives better mixing results than the unsewept raised ramp. Furthermore numerical results show that the swept ramps give better results than the two unswept configurations. introduction Considerable fundamental research has been conducted in response to the increased interests in the development of scramjet propulsion systems. A critical element in the design of the scramjet engine is detailed understanding of the complex flowfield present in different regions of the system over a range of operating conditions. Significant amount of numerical and experimental research have been directed towards injectors design that must produce rapid mixing and combustion of the fuel and air due to the short combustor residence time. In the past years various mixing enhancement schemes including perpendicular and parallel injections have been proposed. One of these schemes is the ramp injection. Critical issues regarding fuel injection and mixing in a scramjet combustor are discussed in detail in Refs. [1-16]. Stouffer et al. [17,18] investigated experimentally both the raised and relieved ramps. The methods that have been 1 Graduate Research Assistant, Department of Mechanical Engineering, Student Member AIAA, ASME. 2 Eminent Professor/Scholar, Department of Mechanical Engineering, Associate Fellow AIAA. 3 Professor and Chair,Mechanical Engineering Department. 4 Associate Professor, Mechanical Engineering Technology Dept, Member ASME, AIAA Copyright © 2000 by the American Institute of Aeronautics and Astronautics, Inc. All rights are reserved. (c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. used to enhance the mixing process for the scramjet combustor are reviewed and summarized by Seiner et al. [19]. The main objective of this research is to study numerically the mixing process in scramjet combustor using unstructured grids. Two different configurations of wall-mounted ramps injectors, raised and relieved ramp injectors, have been investigated.

The RHYFL facility as a high performance expansion tube for scramjet testing

Abstract: While current high enthalpy facilities are capable of generating the correct enthalpies associated with very high speed atmospheric flight, they are incapable of generating the large total pressures needed for full simulation. For accurate aerodynamic, heating and combustion testing in these ground-based facilities, it is desirable to increase their total pressue simulation capabilities. This paper discusses the proposed conversion of RHYFL, which was originally designed to be a shock tube, into the worlds largest expansion tube. If converted to an expansion tube, the RHYFL-X promises to produce high-quality hypersonic flows required for the accurate simulation of atmospheric scramjet flight. Preliminary simulation results are presented and compared to other large testing facilities. These results indicate that the new facility will possess the unique capability of being able to generate test flows that duplicate the free-stream conditions experienced by a scramjet during atmospheric flight.

Reducing skin friction drag

Abstract: The skin friction drag acting on a surface of an article travelling at high speed such as a vehicle at supersonic or particularly hypersonic speed can be reduced by introducing a fuel into the boundary layer under conditions of the fuel introduction to ensure combustion in the boundary layer. The fuel is injected through orifice(s) or a slot provided at the surface so that the fuel enters the passing fluid with a major component of the direction of injection being parallel to the local flow direction. The fuel is injected at supersonic speed, e.g. at a speed of about Mach 1.5 or higher. The invention is applicable to scramjet engines with the fuel being injected around the entire internal circumference of the wall of the scramjet engine, upstream of the commencement of the combustion chamber.

HyShot-T4 Supersonic Combustion Experiments

Abstract: A series of experiments were initiated to investigate the operation of a two-dimensional, hypersonic, airbreathing engine (scramjet) inclined at angles of attack to the freestream. The experiments were undertaken to obtain data for use in the Hyshot flight test program. Experiments on the Hyshot scramjet were under taken in the T4 shock tunnel. Experiments were made at a nominal total enthalpy of 3.0MJkg (exp -1) using a nozzle that produced flows with a Mach number of approximately 6.5. The conditions produced correspond to flight at Mach 7.6 at an altitude range of 35.7-21.4km. A summary of the flow conditions is included. The scramjet was tested at 0, plus 2, plus 4, minus 2 and minus 4 degrees angle of attack. Experiments were also undertaken at 2 and 4 degrees angle of skew.

Mixing Augmentation of Transverse Injection in Scramjet Combustor

Abstract: A method for augmenting the mixing of transverse injection in a scramjet combustor is suggested, and its effectiveness is checked by numerical methods. The intent was to promote streamwise vorticity by modifying the injector geometry using a cavity to increase the mixing rate and penetration. A three-dimensional Navier ‐ Stokes code adopting the upwind method of Edwards’ s low diffusion e ux splitting scheme was used. The k‐! SST turbulence model was used to calculate the turbulent viscosity. The injector models proposed showed remarkable increases of streamwise vorticity and, as a result, signie cant improvements in mixing characteristics such as mixing rate and penetration. However, the proposed models also showed additional losses of stagnation pressure. Themixing characteristicsarestrongly related to the jet-to-cross e ow momentum e ux ratio J. In the case of higher values of J, slower mixing rates, higher penetration, and more losses of stagnation pressure are shown.