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At least 145 records · Page 8

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NASA NTRS1993

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

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

Stouffer, Scott D.↗

NASA NTRS1993

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

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

Baker, N. R.↗

NASA NTRS1993

Numerical study of the performance of swept, curved compression surface scramjet inlets

The paper describes a computational performance-enhancement study in which systematic modifications were applied to a planar-sidewall compression scramjet inlet operating at an entrance Mach number of 4 and at a dynamic pressure of 2040 lbs/sq ft, and the performance of the modified inlet was compared to that of the original geometry by evaluating numerical flowfield solutions. The planar-sidewall compression angle was modified as a function of height using sidewall curvature and simultaneously employing both forward-swept and reverse-swept compression surfaces. The results of calculations illustrate the feasibility of obtaining enhanced performance via contour modifications of a baseline swept-sidewall compression scramjet inlet operating at an entrance Mach number of 4.

Korte, John J.↗

NASA NTRS1993

A numerical study of mixing and combustion in hypervelocity flows through a scramjet combustor model

Interest in high speed, air-breathing propulsion systems such as scramjets has revived in recent years fueled to a large extent by the National Aerospace Plane (NASP) program. These vehicles are expected to fly trans-atmospheric and as a consequence, the Mach number level within the engine/combustor would be rather high (M greater than 5). Ground based testing of such scramjet engines requires a facility that can not only achieve the right Mach number, but also have the proper pressures and temperatures to simulate the combustion processes. At present, only pulse type facilities can provide such high enthalpy flows. The newest of these is the free-piston shock tunnel, T5 located at GALCIT. Recently, a generic combustor model was tested in T5, and the experimental data from that study is analyzed in the present report. The available experimental data from T5 are essentially the static pressures on the injection wall and the one opposite to it. Thus, a principal aim of the present study was to validate the available experimental data by using a proven CFD tool and then investigate the performance characteristics of the combustor model, such as, the mixing efficiency and combustion efficiency. For this purpose, in this study, the code GASP has been used.

Krishamurthy, Ramesh↗

NASA NTRS1995

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

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

Holland, Scott D.↗

NASA NTRS1994

Modeling scramjet combustor flowfields with a grid adaptation scheme

The accurate description of flow features associated with the normal injection of fuel into supersonic primary flows is essential in the design of efficient engines for hypervelocity aerospace vehicles. The flow features in such injections are complex with multiple interactions between shocks and between shocks boundary layers. Numerical studies of perpendicular sonic N2 injection and mixing in a Mach 3.8 scramjet combustor environment are discussed. A dynamic grid adaptation procedure based on the equilibration of spring-mass system is employed to enhanced the description of the complicated flow features. Numerical results are compared with experimental measurements and indicate that the adaptation procedure enhances the capability of the modeling procedure to describe the flow features associated with scramjet combustor components.

Ramakrishnan, R.↗

NASA NTRS1995

Scramjet thrust measurement in a shock tunnel

This note reports tests in a shock tunnel in which a fully integrated scamjet configuration produced net thrust. The experiments not only showed that impluse facilities can be used for assessing thrust performance, but also were a demonstration of the application of a new technique to the measurement of thrust on scramjet configurations in shock tunnels. These two developments are of significance because scramjets are expected to operate at speeds well in excess of 2 km/s, and shock tunnels offer a means of generating high Mach number flows at such speeds.

Paull, A.↗

NASA NTRS1997

Shock Tunnel Studies of Scramjet Phenomena 1995

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

Morgan, R. G.↗

NASA NTRS1997

Shock Tunnel Studies of Scramjet Phenomena 1994

Morgan, R. G.↗

NASA NTRS1998

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

A three-component stress-wave force-balance for a large scramjet has been designed, calibrated and tested in the HYPULSE reflected shock tunnel at GASL Inc., New York. The scramjet model is over 3-foot long and weighs in excess of 90 Ibm. The stress-wave force-balance is comprised of three stress bars which are attached to the model. Calibration results indicate that the force balance responds well within about 1 ms and that the sensitivity of the balance to the distribution of load is not large. Results with and without fuel injection were obtained in the tunnel operated for Mach 7 and Mach 10 flight simulation. These tests showed the force-balance can resolve axial force increments due to combustion of about 40 lb in the presence of model lift forces of 500-700 lb.

Tsai, C.-Y.↗

NASA NTRS2001

MHD Energy Bypass Scramjet Engine

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

Mehta, Unmeel B.↗

NASA NTRS2000

Power Requirement for Nonequilibrium MHD-Bypass Scramjet

It has been suggested previously that the performance of scramjet propulsion system may be improved by the use of magnetohydrodynamic (MHD) energy bypass: an MHD generator could be made to decelerate the flow entering the combustor, thereby improving combustion efficiency, and the electrical power generated could be made to accelerate the flow exiting from the combustor prior to expanding through the nozzle. In one of such proposed schemes, the MHD generator is proposed to be operated at a low temperature and ionization is to be achieved under nonequilibrium by the application of an external power. In the present work, the required power of such an external source is calculated assuming a 100%-efficient nonequilibrium ionization scheme. The power required is that needed to prevent the degree of ionization from reaching equilibrium with the low gas temperature. The flow is seeded with potassium or cesium. Specific impulse is calculated with and without turbulent friction. The results show that, for typical intended flight conditions, the specific impulse obtained is substantially higher than that of a typical scramjet, but the required external-power is several times that of the power generated in the MHD generator.

Park, Chul↗

NASA NTRS2000

MHD Energy Bypass Scramjet Performance with Real Gas Effects

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.

Park, Chul↗

NASA NTRS2003

Fuel Injector Design Optimization for an Annular Scramjet Geometry

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

Steffen, Christopher J., Jr.↗

NASA NTRS2004

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

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

Mohieldin, T. O.↗

NASA NTRS2005

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

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

Marshall, Laurie A.↗

NASA NTRS2005

Shock Interaction Control for Scramjet Cowl Leading Edges

An experimental study was conducted to qualitatively determine the effectiveness of stagnation-region gas injection in protecting a scramjet cowl leading edge from the intense heating produced by Type III and Type IV shock interactions. The model consisted of a two-dimensional leading edge, representative of that of a scramjet cowl. Tests were conducted at a nominal freestream Mach number of 6. Gaseous nitrogen was supersonically injected through the leading-edge nozzles at various mass flux ratios and with the model pitched at angles of 0deg and -20deg relative to the freestream flow. Qualitative data, in the form of focusing and conventional schlieren images, were obtained of the shock interaction patterns. Results indicate that large shock displacements can be achieved and both the Type III and IV interactions can be altered such that the interaction does not impinge on the leading edge surface.

Albertson, Cindy W.↗

NASA NTRS2005

Hyper-X Mach 7 Scramjet Design, Ground Test and Flight Results

The successful Mach 7 flight test of the Hyper-X (X-43) research vehicle has provided the major, essential demonstration of the capability of the airframe integrated scramjet engine. This flight was a crucial first step toward realizing the potential for airbreathing hypersonic propulsion for application to space launch vehicles. However, it is not sufficient to have just achieved a successful flight. The more useful knowledge gained from the flight is how well the prediction methods matched the actual test results in order to have confidence that these methods can be applied to the design of other scramjet engines and powered vehicles. The propulsion predictions for the Mach 7 flight test were calculated using the computer code, SRGULL, with input from computational fluid dynamics (CFD) and wind tunnel tests. This paper will discuss the evolution of the Mach 7 Hyper-X engine, ground wind tunnel experiments, propulsion prediction methodology, flight results and validation of design methods.

Ferlemann, Shelly M.↗