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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: 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: Publisher: ISBN: Category : Languages : en Pages : 152
Book Description
This report investigated fundamental aerothermal phenomena in hypersonic flow, with particular emphasis on viscous/inviscid interaction phenomena. The experimental studies were conducted to examine the changes in the structure at the base of a hypersonic turbulent boundary layer as it is subjected to a strong self-induced pressure gradient in regions of shock wave/ boundary layer interaction. The initial phase of the theoretical program was directed toward summarizing existing techniques for obtaining Navier/Stokes solutions for laminar flow over flat plates in hypersonic flow. In the experimental program, surface and flow field measurements were made to examine the detailed flow mechanics associated with turbulent boundary layer separation over a large cone flare model at Mach numbers 11, 13 and 16 and Reynolds numbers up to 100 million. Holography measurements were used to examine the flow field regions of hypersonic shock wave/turbulent boundary layer interaction. In this preliminary investigation of the application of holography in a 96 inch Shock Tunnel, holographic interferograms were obtained for viscous/inviscid interactions at Mach numbers 11 and 13 and Reynolds numbers up to 30 million, nominally. Flow field studies were made for flat plate/wedge, cone/flare and incident shock configurations. The quantitative results presented offer new information as well as a potential to obtain density measurements in other types of hypersonic flow; however, they also reveal important concerns which need to be resolved before the interferometric data can be claimed to give accurate measurements of these flows. Two appendices include an invited review paper on aerothermal problems associated with hypersonic flight.
Author: William H. Heiser Publisher: AIAA ISBN: 9781563470356 Category : Science Languages : en Pages : 632
Book Description
An almost entirely self-contained engineering textbook primarily for use in undergraduate and graduate courses in airbreathing propulsion. It provides a broad and basic introduction to the elements needed to work in the field as it develops and grows. Homework problems are provided for almost every individual subject. An extensive array of PC-based user-friendly computer programs is provided in order to facilitate repetitious and/or complex calculations. Annotation copyright by Book News, Inc., Portland, OR
Author: Konstantinos Kontis Publisher: Springer ISBN: 3319731807 Category : Technology & Engineering Languages : en Pages : 408
Book Description
This edited monograph contains the proceedings of the International Shock Interaction Symposium, which emerged as an heir to both the Mach Reflection and Shock Vortex Interaction Symposia. These scientific biannual meetings provide an ideal platform to expose new developments and discuss recent challenges in the field of shock wave interaction phenomena. The goal of the symposia is to offer a forum for international interaction between young and established scientists in the field of shock and blast wave interaction phenomena. The target audience of this book comprises primarily researchers and experts in the field of shock waves, but the book may also be beneficial for young scientists and graduate students alike.
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 Administration (NASA) Publisher: Createspace Independent Publishing Platform ISBN: 9781723575044 Category : Languages : en Pages : 234
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...
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Author: Davis H. Crawford
Author: Herbert S. Glick
Author: William H. Heiser
Author: Michael S. Holden
Author: Michael S. Holden
Author: Konstantinos Kontis
Author: National Aeronautics and Space Adm Nasa
Author: National Aeronautics and Space Administration (NASA)