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Author: C. Herbert Law Publisher: ISBN: Category : Aerodynamic heating Languages : en Pages : 52
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: C. Herbert Law Publisher: ISBN: Category : Aerodynamic heating Languages : en Pages : 52
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
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: Seymour M. Bogdonoff Publisher: ISBN: Category : Languages : en Pages : 75
Book Description
An extensive experimental study of three-dimensional shock wave turbulent boundary layer interactions caused by shock generators defined solely by angles has been carried out at Mach 3. Sharp fins, sharp swept fins, swept wedges, and semi-cones have been used to generate a wide range of shock waves. The interaction of these waves with turbulent boundary layers has been investigated by surface flow visualization, mean surface static pressure distributions, flowfield surveys of total pressure and yaw, and several flowfield visualization techniques. Some exploratory high frequency surface pressure measurements have been carried out to evaluate the steadiness of these interactions. Scaling laws for both surface and flowfield features have been derived. Some limited studies were carried out at a Mach number of 2. A flowfield study has shown that the initial part of interactions caused by the same strength and geometrical shock wave generated by different shock generators are all similar. The 'footprints' of the interactions, as shown by surface flow visualization, can be categorized as approximately conical or cylindrical, and the boundaries between these two regions have been defined for both Mach 3 and Mach 2. There are still questions with regards to the detailed flowfield structures and physical mechanisms, but the three-dimensional interactions appeared to be less unsteady than that of two-dimensional separated flows.
Author: National Aeronautics and Space Administration (NASA) Publisher: Createspace Independent Publishing Platform ISBN: 9781722602345 Category : Languages : en Pages : 52
Book Description
Experimental data for two three-dimensional intersecting shock-wave/turbulent boundary-layer interaction flows at Mach 8.3 are presented. The test bodies, composed of two sharp fins fastened to a flat-plate test bed, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface pressure and heat transfer distributions as well as mean flow-field surveys both in the undisturbed and interaction regimes. The data are presented in a convenient form to be used to validate existing or future computational models of these hypersonic flows. The data are also on a 3.5-inch diskette included and are available through E-mail. Kussoy, M. I. and Horstman, K. C. Ames Research Center...
Author: Doyle D. Knight Publisher: ISBN: Category : Languages : en Pages : 50
Book Description
The focus of the research effort is the understanding of three-dimensional shock wave-turbulent boundary layer interactions. The approach uses the full mean compressible Navier-Stokes equations with turbulence incorporated through the algebraic turbulent eddy viscosity model of Baldwin and Lomax. During the present year of the research effort, the three-dimensional shock boundary layer interaction generated by a 10 deg sharp fin has been computed at Mach 3 for a Reynolds number 280000. These results, together with previous computations of the same configuration at Reynolds number = 930000, are compared with experimental data for pitot pressure and yaw angle. The agreement with the experimental data is good, and the theory accurately predicts the recovery of the boundary layer downstream of the interaction of Reynolds number = 280000. The computed flowfield is employed to analyze the structure of the 3-D interaction through contour plots of flow variables. Also, during the present year, the investigation of the 2-D turbulent supersonic compression corner at Mach 3 was completed. The relaxation modification to the Baldwin-Lomax model was found to yield reasonably accurate predictions of the upstream propagation of the surface for the Reynolds number range investigated. An additional computation at Mach 2 was performed, and the results were in general in agreement with the previous conclusions. (Author).
Author: Publisher: ISBN: Category : Languages : en Pages : 71
Book Description
The focus of the research effort is the understanding of three-dimensional shock wave-turbulent boundary layer interactions. The approach uses the full mean compressible Navier-Stokes equations with turbulence incorporated through the algebraic turbulent eddy viscosity model of Baldwin and Lomax. During the present year of the research effort, the three-dimensional shock boundary layer interaction generated by a 10 deg sharp fin has been computed at Mach 3 for a Reynolds number 280000. These results, together with previous computations of the same configuration at Reynolds number = 930000, are compared with experimental data for pitot pressure and yaw angle. The agreement with the experimental data is good, and the theory accurately predicts the recovery of the boundary layer downstream of the interaction of Reynolds number = 280000. The computed flowfield is employed to analyze the structure of the 3-D interaction through contour plots of flow variables. Also, during the present year, the investigation of the 2-D turbulent supersonic compression corner at Mach 3 was completed. The relaxation modification to the Baldwin-Lomax model was found to yield reasonably accurate predictions of the upstream propagation of the surface for the Reynolds number range investigated. An additional computation at Mach 2 was performed, and the results were in general in agreement with the previous conclusions. (Author).
Author: C. Herbert Law
Author: C. Herbert Law
Author: C. Herbert Law
Author: Doyle D. Knight
Author: B. Shapey
Author: Holger Babinsky
Author: Seymour M. Bogdonoff
Author: Morgan Lee Funderburk
Author: National Aeronautics and Space Administration (NASA)
Author: Doyle D. Knight
Author: