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.

Laminar and turbulent boundary layer interactions with externally generated shock waves.

Tabulated data from a series of experimental studies of the interaction of a shock wave with a turbulent boundary layer in axisymmetric flow configurations is presented. The studies were conducted at the walls of circular wind tunnels and on the cylindrical centerbody of an annular flow channel. Detailed pitot pressure profiles and wall static pressure profiles upstream of, within and downstream of the interaction region are given. Results are presented for flows at nominal freestream Mach Numbers of 2, 3 and 4. For studies at the tunnel sidewalls, the shock waves were produced by conical shock generators mounted on the centerline of the wind tunnel at zero angle of attack. The annular ring generator was used to produce the shock wave at the centerbody of the annular flow channel. The effects of boundary layer bleed were examined in the investigation. Both bleed rate and bleed location were studied. Most of the bleed studies were conducted with bleed holes drilled normal to the wall surface but the effects of slot suction were also examined. A summary of the principal results and conclusions is given.

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.