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Author: National Aeronautics and Space Adm Nasa Publisher: ISBN: 9781730766190 Category : Languages : en Pages : 66
Book Description
Two classes of shock-wave boundary-layer interactions were studied experimentally in a shock tunnel in which a low Reynolds number, turbulent flow at Mach 8 was developed on a cold, flat test surface. The two classes of interactions were: (1) a swept interaction generated by a wedge ('fin') mounted perpendicularly on the flat plate; and (2) a two-dimensional, unseparated interaction induced by a shock impinging near an expansion corner. The swept interaction, with wedge angles of 5-20 degrees, was separated and there was also indication that the strongest interactions prossessed secondary separation zones. The interaction spread out extensively from the inviscid shock location although no indication of quasi-conical symmetry was evident. The surface pressure from the upstream influence to the inviscid shock was relatively low compared to the inviscid downstream value but it rose rapidly past the inviscid shock location. However, the surface pressure did not reach the downstream inviscid value and reasons were proposed for this anomalous behavior compared to strongly separated, supersonic interactions. The second class of interactions involved weak shocks impinging near small expansion corners. As a prelude to studying this interaction, a hypersonic similarity parameter was identified for the pure, expansion corner flow. The expansion corner severely damped out surface pressure fluctuations. When a shock impinged upstream of the corner, no significant changes to the surface pressure were found as compared to the case when the shock impinged on a flat plate. But, when the shock impinged downstream of the corner, a close coupling existed between the two wave systems, unlike the supersonic case. This close coupling modified the upstream influence. Regardless of whether the shock impinged ahead or behind the corner, the downstream region was affected by the close coupling between the shock and the expansion. Not only was the mean pressure distribution modified but the un...
Author: National Aeronautics and Space Adm Nasa Publisher: ISBN: 9781730766190 Category : Languages : en Pages : 66
Book Description
Two classes of shock-wave boundary-layer interactions were studied experimentally in a shock tunnel in which a low Reynolds number, turbulent flow at Mach 8 was developed on a cold, flat test surface. The two classes of interactions were: (1) a swept interaction generated by a wedge ('fin') mounted perpendicularly on the flat plate; and (2) a two-dimensional, unseparated interaction induced by a shock impinging near an expansion corner. The swept interaction, with wedge angles of 5-20 degrees, was separated and there was also indication that the strongest interactions prossessed secondary separation zones. The interaction spread out extensively from the inviscid shock location although no indication of quasi-conical symmetry was evident. The surface pressure from the upstream influence to the inviscid shock was relatively low compared to the inviscid downstream value but it rose rapidly past the inviscid shock location. However, the surface pressure did not reach the downstream inviscid value and reasons were proposed for this anomalous behavior compared to strongly separated, supersonic interactions. The second class of interactions involved weak shocks impinging near small expansion corners. As a prelude to studying this interaction, a hypersonic similarity parameter was identified for the pure, expansion corner flow. The expansion corner severely damped out surface pressure fluctuations. When a shock impinged upstream of the corner, no significant changes to the surface pressure were found as compared to the case when the shock impinged on a flat plate. But, when the shock impinged downstream of the corner, a close coupling existed between the two wave systems, unlike the supersonic case. This close coupling modified the upstream influence. Regardless of whether the shock impinged ahead or behind the corner, the downstream region was affected by the close coupling between the shock and the expansion. Not only was the mean pressure distribution modified but the un...
Author: Holger Babinsky Publisher: Cambridge University Press ISBN: 1139498649 Category : Technology & Engineering Languages : en Pages : 481
Book Description
Shock wave-boundary-layer interaction (SBLI) is a fundamental phenomenon in gas dynamics that is observed in many practical situations, ranging from transonic aircraft wings to hypersonic vehicles and engines. SBLIs have the potential to pose serious problems in a flowfield; hence they often prove to be a critical - or even design limiting - issue for many aerospace applications. This is the first book devoted solely to a comprehensive, state-of-the-art explanation of this phenomenon. It includes a description of the basic fluid mechanics of SBLIs plus contributions from leading international experts who share their insight into their physics and the impact they have in practical flow situations. This book is for practitioners and graduate students in aerodynamics who wish to familiarize themselves with all aspects of SBLI flows. It is a valuable resource for specialists because it compiles experimental, computational and theoretical knowledge in one place.
Author: Michael S. Holden Publisher: ISBN: Category : Languages : en Pages : 76
Book Description
A detailed description is given of the theoretical and experimental studies of shock wave-laminar boundary layer interaction on curved compression surfaces in hypersonic flow. Integral forms of the boundary layer equations were used for the conservation of mass, streamwise momentum, normal momentum, moment of streamwise momentum and energy to describe the development of attached and separated boundary layers on a curve surface. In the experimental study, skin friction, heat transfer and pressure measurements were made on a series of flat plate-cylindrical arc-wedge compression surfaces. Pitot and cone pressure measurements were made above the surface in the separation and reattachment regions of separated flows to estimate the static pressure difference across the boundary layer. The radius of curvature of the cylindrical arc, the inclination of the wedge, and the unit Reynolds number of the free-stream were varied to examine their effect on the properties of both attached and separated interaction regions over the models. (Author).
Author: National Aeronautics and Space Administration (NASA) Publisher: Createspace Independent Publishing Platform ISBN: 9781723575211 Category : Languages : en Pages : 238
Book Description
Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident-shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not sig...
Author: Ozer Igra Publisher: Springer ISBN: 3319237454 Category : Technology & Engineering Languages : en Pages : 480
Book Description
This comprehensive and carefully edited volume presents a variety of experimental methods used in Shock Waves research. In 14 self contained chapters this 9th volume of the “Shock Wave Science and Technology Reference Library” presents the experimental methods used in Shock Tubes, Shock Tunnels and Expansion Tubes facilities. Also described is their set-up and operation. The uses of an arc heated wind tunnel and a gun tunnel are also contained in this volume. Whenever possible, in addition to the technical description some typical scientific results obtained using such facilities are described. Additionally, this authoritative book includes techniques for measuring physical properties of blast waves and laser generated shock waves. Information about active shock wave laboratories at different locations around the world that are not described in the chapters herein is given in the Appendix, making this book useful for every researcher involved in shock/blast wave phenomena.
Author: Publisher: ISBN: Category : Languages : en Pages : 0
Book Description
Experimental studies were conducted in conjunction with computations in a code validation exercise to examine the ability of DSMC and Navier-Stokes techniques to predict the complex characteristics of regions of shock/shock and shock/boundary layer interactions in hypervelocity flows. In the experimental program, detailed heat transfer and pressure measurements in laminar regions of shock wave/boundary layer interaction, and shock/shock interaction, over hollow cylinder/flare and double cone configurations in hypersonic flow. In the best Navier-Stokes solutions the structure and density of the flowfield was captured exactly over both the hollow cylinder/flare and double cone models. The detailed characteristics of the distribution of pressure and heating through the interaction regions were well predicted. In general, for these relatively high-density flows, the DSMC solutions did not capture the characteristics of well-separated flows. In Part 2 of this program, experimental studies have been conducted in the LENS I shock tunnel to investigate the characteristics and performance of full-scale scramjet engines at fully duplicated flight conditions for a range of dynamic pressures at Mach 7. In this program we obtained detailed heat transfer and pressure measurements together with high speed Schlieren and infrared imagery to determine the characteristics of a generic engine, the flow path of which was configured to be similar to the Air Force Hy-Tech configuration for studies with hydrocarbon and hydrogen fuels.
Author: National Aeronautics and Space Adm Nasa
Author: National Aeronautics and Space Adm Nasa
Author: Holger Babinsky
Author: Michael S. Holden
Author: R. J. Hawboldt
Author: National Aeronautics and Space Administration (NASA)
Author: Michael S. Holden
Author: 
Author: Ozer Igra
Author: 
Author: James M. Kendall