Interaction of Two Glancing, Crossing Shock Waves with a Turbulent Boundary-Layer at Various Mach Numbers PDF Download
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Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781722499792
Category :
Languages : en
Pages : 98
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Book Description
A preliminary experimental investigation was conducted to study two crossing, glancing shock waves of equal strengths, interacting with the boundary-layer developed on a supersonic wind tunnel wall. This study was performed at several Mach numbers between 2.5 and 4.0. The shock waves were created by fins (shock generators), spanning the tunnel test section, that were set at angles varying from 4 to 12 degrees. The data acquired are wall static pressure measurements, and qualitative information in the form of oil flow and schlieren visualizations. The principle aim is two-fold. First, a fundamental understanding of the physics underlying this flow phenomena is desired. Also, a comprehensive data set is needed for computational fluid dynamic code validation. Results indicate that for small shock generator angles, the boundary-layer remains attached throughout the flow field. However, with increasing shock strengths (increasing generator angles), boundary layer separation does occur and becomes progressively more severe as the generator angles are increased further. The location of the separation, which starts well downstream of the shock crossing point, moves upstream as shock strengths are increased. At the highest generator angles, the separation appears to begin coincident with the generator leading edges and engulfs most of the area between the generators. This phenomena occurs very near the 'unstart' limit for the generators. The wall pressures at the lower generator angles are nominally consistent with the flow geometries (i.e. shock patterns) although significantly affected by the boundary-layer upstream influence. As separation occurs, the wall pressures exhibit a gradient that is mainly axial in direction in the vicinity of the separation. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout. Hingst, Warren R. and Williams, Kevin E. Glenn Research Center...
Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781722499792
Category :
Languages : en
Pages : 98
Get Book Here
Book Description
A preliminary experimental investigation was conducted to study two crossing, glancing shock waves of equal strengths, interacting with the boundary-layer developed on a supersonic wind tunnel wall. This study was performed at several Mach numbers between 2.5 and 4.0. The shock waves were created by fins (shock generators), spanning the tunnel test section, that were set at angles varying from 4 to 12 degrees. The data acquired are wall static pressure measurements, and qualitative information in the form of oil flow and schlieren visualizations. The principle aim is two-fold. First, a fundamental understanding of the physics underlying this flow phenomena is desired. Also, a comprehensive data set is needed for computational fluid dynamic code validation. Results indicate that for small shock generator angles, the boundary-layer remains attached throughout the flow field. However, with increasing shock strengths (increasing generator angles), boundary layer separation does occur and becomes progressively more severe as the generator angles are increased further. The location of the separation, which starts well downstream of the shock crossing point, moves upstream as shock strengths are increased. At the highest generator angles, the separation appears to begin coincident with the generator leading edges and engulfs most of the area between the generators. This phenomena occurs very near the 'unstart' limit for the generators. The wall pressures at the lower generator angles are nominally consistent with the flow geometries (i.e. shock patterns) although significantly affected by the boundary-layer upstream influence. As separation occurs, the wall pressures exhibit a gradient that is mainly axial in direction in the vicinity of the separation. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout. Hingst, Warren R. and Williams, Kevin E. Glenn Research Center...
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 100
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Book Description
Author: Warren R. Hingst
Publisher:
ISBN:
Category : Shock waves
Languages : en
Pages : 93
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Book Description
Author: C. Herbert Law
Publisher:
ISBN:
Category : Aerodynamic heating
Languages : en
Pages : 52
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Book Description
Experimental results of an investigation of the three-dimensional interaction between a skewed shock wave and a turbulent boundary layer are presented. Surface pressure and heat transfer distributions and oil flow photographs were obtained at a freestream Mach number of 5.85 and two Reynolds numbers of ten and twenty million per foot. The model configuration consisted of a shock generator mounted perpendicularly to a flat plate. The shock generator leading edge was sharp and nonswept and intersected the flat plate surface about 8.5 inches downstream of the flat plate leading edge. The shock generator surface was 7.55 inches long and 3 inches high and its angle to the freestream flow was adjusted from 4 to 20 degrees. The generated shock waves were of sufficient strength to produce turbulent boundary layer separation on the flat plate surface.
Author: Holger Babinsky
Publisher: Cambridge University Press
ISBN: 1139498649
Category : Technology & Engineering
Languages : en
Pages : 481
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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:
Publisher:
ISBN:
Category : Aeronautics
Languages : en
Pages : 584
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Book Description
Author: Hyun D. Kim
Publisher:
ISBN:
Category :
Languages : en
Pages : 14
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Book Description
Author: C. Herbert Law
Publisher:
ISBN:
Category : Aerodynamic load
Languages : en
Pages : 120
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Book Description
Author: Konstantinos Kontis
Publisher: Springer Science & Business Media
ISBN: 3642256856
Category : Science
Languages : en
Pages : 1122
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Book Description
The University of Manchester hosted the 28th International Symposium on Shock Waves between 17 and 22 July 2011. The International Symposium on Shock Waves first took place in 1957 in Boston and has since become an internationally acclaimed series of meetings for the wider Shock Wave Community. The ISSW28 focused on the following areas: Blast Waves, Chemically Reacting Flows, Dense Gases and Rarefied Flows, Detonation and Combustion, Diagnostics, Facilities, Flow Visualisation, Hypersonic Flow, Ignition, Impact and Compaction, Multiphase Flow, Nozzle Flow, Numerical Methods, Propulsion, Richtmyer-Meshkov, Shockwave Boundary Layer Interaction, Shock Propagation and Reflection, Shock Vortex Interaction, Shockwave Phenomena and Applications, as well as Medical and Biological Applications. The two Volumes contain the papers presented at the symposium and serve as a reference for the participants of the ISSW 28 and individuals interested in these fields.
Author: H. Kubota
Publisher:
ISBN:
Category :
Languages : en
Pages : 243
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Book Description
Three-dimensional glancing-shock/turbulent boundary-layer interaction has been investigated at the Cranfield Institute of Technology in two separate test programmes using a 2.5 x 2.5 inch intermittent tunnel and a 9 x 9 inch continuous tunnel, at a Mach number of approximately 2.5. The experimental results include oil-flow pictures, vapour-screen photographs, surface static pressure distributions, local heat transfers, liquid crystal pictures of surface temperature, and viscous layer surveys. The test data indicate that the interaction region consists of two different viscous flows, the side-wall boundary layer and an induced layer originating near the shock generator root and crossing the path of the side-wall boundary layer. In this flow field model, no flow separation appears as long as the surface stream lines of the side-wall boundary layer can be pliable enough to be bent along the edge of the induced layer, even when the surface-flow deflection exceeds the shock angle. However, an ordinary separation does take place when the induced layer forces the surface stream lines to deflect beyond a maximum permissible angle. (Author).
