Separation measurements of supersonic turbulent boundary layers over compression corners
Separation measurements of supersonic turbulent boundary layers over compression corners

Author: Walter Bruce Gillette

Publisher:

Published: 1967

Total Pages: 111

ISBN-13:

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To the designer of modern high-speed aerospace vehicles, separation of the boundary layer before macroscopic surface features is an important problem. A typical feature of interest is a compression corner, or ramp. To aid in the development of comprehensive theories concerning boundary layer separation before a compression corner, a series of experimental investigations were conducted for the adiabatic flow of a supersonic compressible gas over a compression corner. Tests were conducted at free-stream Mach Numbers of 2.00 to 4.00 in increments of 0.50, at flat plate Reynolds Numbers of 0.75 x 10 to the 7th and 1.5 x 10 to the 7th, and for compression angles of 10.3 deg, 20.1 deg, 30.5 deg, and 39.9 deg. Static pressure surveys of the flow ahead of and over the compression corner were made. These measurements were supplemented by high-speed schlieren photographs and shadowgraphs. The separation of the turbulent compressible boundary layer was found to have strong dependence on both the Mach Number and the Reynolds Number. For Mach Numbers less than 3.00, the separation distance ahead of the compression corner decreased with increasing Mach Number. For Mach Numbers of 3.50 and 4.00, the separation distance increased with Mach Number. At all Mach Numbers, an increase in Reynolds Number increased the separation distance. The Reynolds Number influence was greater at the higher Reynolds Numbers. Unsteadiness in the separation geometry occurred for separation distances greater than six or eight boundary layer thicknesses. The separation was found to result from a free interaction of the flow phenomena involved. (Author).

Turbulent Shear Layers in Supersonic Flow
Turbulent Shear Layers in Supersonic Flow

Author: Alexander J. Smits

Publisher: Springer Science & Business Media

Published: 2006-05-11

Total Pages: 418

ISBN-13: 0387263055

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A good understanding of turbulent compressible flows is essential to the design and operation of high-speed vehicles. Such flows occur, for example, in the external flow over the surfaces of supersonic aircraft, and in the internal flow through the engines. Our ability to predict the aerodynamic lift, drag, propulsion and maneuverability of high-speed vehicles is crucially dependent on our knowledge of turbulent shear layers, and our understanding of their behavior in the presence of shock waves and regions of changing pressure. Turbulent Shear Layers in Supersonic Flow provides a comprehensive introduction to the field, and helps provide a basis for future work in this area. Wherever possible we use the available experimental work, and the results from numerical simulations to illustrate and develop a physical understanding of turbulent compressible flows.

Two-Dimensional Compression Corner and Planar Shock Wave Interactions with a Supersonic, Turbulent Boundary Layer
Two-Dimensional Compression Corner and Planar Shock Wave Interactions with a Supersonic, Turbulent Boundary Layer

Author: C. Herbert Law

Publisher:

Published: 1975

Total Pages: 118

ISBN-13:

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Experimental data have been obtained to describe the interactions between a turbulent boundary layer and (1) a two-dimensional compression corner and (2) an externally generated planar shock wave. Investigations were conducted at Mach number 3 over a range of Reynolds numbers from 10 to 100 million under adiabatic wall conditions. The effects of compression corner angle, planar shock wave strength, and Reynolds number on the length of separation and upstream influence were obtained. The incipient separation conditions and three-dimensional effects were also investigated. The separation and upstream influence lengths were found to increase with increasing Reynolds number for fixed overall pressure rise. The overall pressure rise for incipient separation is approximately the same for the compression corner interaction and the planar shock wave interaction, and increases with increasing Reynolds number. Turbulent boundary layer separation was found to be of the free interaction type, whereby the separation angle and the pressure distribution through separation were found to be independent of Reynolds number, overall pressure rise, and type of disturbance. For the case of the externally generated shock wave, the span of the shock generator had to be reduced in order to eliminate an initially significant influence of the sidewall shock wave-boundary layer interaction on the test region. (Author).

Two-Dimensional Compression Corner and Planar Shock Wave Interactions with a Supersonic, Turbulent Boundary Layer
Two-Dimensional Compression Corner and Planar Shock Wave Interactions with a Supersonic, Turbulent Boundary Layer

Author:

Publisher:

Published: 1975

Total Pages: 0

ISBN-13:

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Experimental data have been obtained to describe the interactions between a turbulent boundary layer and (1) a two-dimensional compression corner and (2) an externally generated planar shock wave. Investigations were conducted at Mach number 3 over a range of Reynolds numbers from 10 to 100 million under adiabatic wall conditions. The effects of compression corner angle, planar shock wave strength, and Reynolds number on the length of separation and upstream influence were obtained. The incipient separation conditions and three-dimensional effects were also investigated. The separation and upstream influence lengths were found to increase with increasing Reynolds number for fixed overall pressure rise. The overall pressure rise for incipient separation is approximately the same for the compression corner interaction and the planar shock wave interaction, and increases with increasing Reynolds number. Turbulent boundary layer separation was found to be of the free interaction type, whereby the separation angle and the pressure distribution through separation were found to be independent of Reynolds number, overall pressure rise, and type of disturbance. For the case of the externally generated shock wave, the span of the shock generator had to be reduced in order to eliminate an initially significant influence of the sidewall shock wave-boundary layer interaction on the test region. (Author).

Experimental Study on High Subsonic Turbulent Flow Incipient Separation
Experimental Study on High Subsonic Turbulent Flow Incipient Separation

Author: Jain-Ming Wu

Publisher:

Published: 1976

Total Pages: 68

ISBN-13:

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For flow over a two-dimensional ramp compression corner case, an experimental investigation to determine the incipient separation was carried out at Mach numbers between 0.55 and 0.9 and Reynolds numbers (based on undisturbed boundary layer thickness) between 350,000 and 690,000. Detailed surface pressure, pitot traversing and oil flow data were obtained for each ramp angle case. Two-dimensionality of the ramp compression corners was verified by the surface oil flow. A major finding of this study is that the incipient separation ramp angle is relatively independent of Mach number and Reynolds number within the range studied. The incipient separation ramp angle was found to be about 22.5 degrees.

Research on Supersonic Turbulent Separated and Reattached Flows
Research on Supersonic Turbulent Separated and Reattached Flows

Author: Seymour M. Bogdonoff

Publisher:

Published: 1975

Total Pages: 94

ISBN-13:

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Basic research programs are reported with fundamental applications to supersonic flight. The experimental studies made use of the unique capabilities of the high Reynolds number Mach 3 facility. The experimental programs concentrated on phenomena associated with incipient separation and separation of turbulent boundary layers over a large Reynolds number range. The reattachment phenomena of a shear layer was also studied in great depth.

Shock Wave-Boundary-Layer Interactions
Shock Wave-Boundary-Layer Interactions

Author: Holger Babinsky

Publisher: Cambridge University Press

Published: 2011-09-12

Total Pages: 481

ISBN-13: 1139498649

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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.