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Aerospace Sciences Meeting and Exhibit, 35th, Reno, NV, Jan. 6-9, 1997, 資料番号: AC1000072000, レポート番号: AIAA-1997-0292

In the present study supersonic mixing flows with parallel injectors are investigated experimentally and computationally. In the experiments the flow fields are visualized by the Schlieren method and flow technique. Also surface pressure distributions in the whole region are measured and complicated flow structure is revealed. In the present study three-dimensional full Navier-Stokes equations are solved for simulating threedimensional supersonic mixing flow problems in detail.The results show the present numerical code could capture the flow structure quantitatively to some extent. Also the results suggest the detailed and reliable experiments can be quite important in set for * Associate Professor, Senior Member, AIAA ** Research Scientists, Research Institute *** Graduate Students **** Research Scientist Copyright © 1997 by S. ASO, Y. YAMANE, K. UMII, K. TOKUNAGA, Y. ANDO and K. SAKATA. Published by The American Institute of Aeronautics and Astronautics, Inc., with permission. the code validation of numerical simulations.

A subscale scramjet engine was designed on the basis of the component study results. Both a heat-sink model and a water-cooled model were fabricated. The heat-sink engine has been tested in a free-jet engine test facility simulating flight conditions of Mach 4 at 20 km altitude and Mach 6 at 25 km. In the Mach 4 test, successful ignition of the main fuel was achieved by the plasma torch igniters with pilot fuel injection. However, in the Mach 6 test, only partial ignition of the fuel was achieved with the same configuration. Wall pressure and temperature distributions as well as forces were measured.

A method for identifying and evaluating the thrust losses in an aerospace engine is first applied to one-dimensional scramjet engine flows with coupled loss mechanisms. Overall engine thrust losses are related directly to local increases in irreversible entropy due to friction, heat transfer, mixing, nonequilibrium reaction, and shocks. This method is then extended to enable the evaluation of thrust losses in multidimensional flows. The fundamental relationship between performance assessment utilizing multidimensional flow fields and cycle analysis performance prediction is shown and discussed. (Author)

Aerospace Sciences Meeting and Exhibit, 34th, Reno, NV, Jan. 15-18, 1996, 資料番号: AC1000071000, レポート番号: AIAA-1996-0380

The results of a feasibility study of a hypersonic waverider research vehicle with a hydrocarbon scramjet engine are presented. The integrated waverider/scramjet geometry is first optimized with a vehicle synthesis code to produce a maximum product of the lift-to-drag ratio and the cycle specific impulse, hence cruise range. Computational fluid dynamics (CFD) is then employed to provide a nose-to-tail analysis of the system at the on-design conditions. Some differences are noted between the results of the two analysis techniques. A comparison of experimental, engineering analysis and CFD results on a waverider forebody are also included for validation.

31st Aerospace Sciences Meeting & Exhibit, Reno, NV, January 11-14, 1993, 資料番号: AC1000069000, レポート番号: AIAA-1993-0739

The Institute for Computer Applications in Science and Engineering (ICASE) and NASA Langley Research Center (LaRC) hosted a workshop on October 2--4, 1989 to discuss some combustion problems of technological interest to LaRC and to foster interaction with the academic community in these research areas. The topics chosen for this purpose were flame structure, flame holding/extinction, chemical kinetics, turbulence-kinetics interaction, transition to detonation, and reacting free shear layers. This document contains the papers and edited versions of general discussions on these topics. The lead paper set the stage for the meeting by discussing the status and issues of supersonic combustion relevant to the scramjet engine. Experts were then called upon to review the current knowledge in the aforementioned areas, to focus on how this knowledge can be extended and applied to high-speed combustion, and to suggest future directions of research in these areas.

This paper is devoted to the analyses and optimization of simple and sophisticated cycles, particularly for various gas turbine engines and aero-engines (including scramjet engine) to achive the maximum performance. The optimization of such criteria as thermal efficiency, specific output and total performance for gas turbine engines, and overall efficiency, non-dimensional thrust and specific impulse for aero-engines have been performed by the optimization procedure with multiplier method. The comparisons of results with analytical solutions establishes the validity of the optimization procedure.Copyright © 1990 by ASME

Although the concept of using a supersonic combustion ramjet engine to propel a vehicle at hypersonic speeds through the atmosphere has been thoroughly established over the past two decades, and analyses for determining the performance of each of its components documented, no complete engine cycle analysis has been presented to date. The purpose of this paper is to review and discuss the individual component analyses previously developed, combine them into a unified cycle analysis, and present the methodology needed to optimize the design of a scramjet engine using the unified cycle analysis. A specific example applying this optimization procedure to a missile design is also included.

Fuel injection and mixing and piloted ignition in supersonic flow design principles for scramjet engine, evaluating eddy viscosity model for parallel injection

Review of the results of an experimental investigation of the performance of an inlet for an integrated scramjet engine concept at Mach 6. Following a description of the inlet design and test model, the Mach 6 experimental results obtained are presented in terms of integrated performance parameters.