Results of a study of hypersonic-turbulent boundary-layer separation over axisymmetric bodies are presented. Pressure, heat-transfer and flow visualization data were obtained for a cone-cylinder-flare configuration. The experimental data were correlated with a theoretical prediction. (Author).

Results of a study of hypersonic-turbulent boundary-layer separation over axisymmetric bodies are presented. Pressure, heat-transfer and flow visualization data were obtained for a cone-cylinder-flare configuration. The experimental data were correlated with a theoretical prediction. (Author).

Experiments at supersonic speeds and at Mach 8 were conducted to determine the conditions which govern the extent of shock-induced laminar flow separations on axisymmetric configurations at zero yaw and without heat transfer. From an extensive correlation of surface pressure data and schlieren photographs, it is shown that the extent of reverse flow is essentially a function of the ratio of the wetted length to the flare divided by the laminar boundary thickness there. As a result, the relative extent of laminar flow separation decreases with a unit Reynolds number increase and grows through an increase in Mach number. Finally, increasing the flare angle increases the length of the reverse flow region.

This is a survey paper on the subject of hypersonic boundary-layer transition. Part 1 discusses boundary-layer stability theory, hypersonic boundary-layer stability experiments, and a comparison between theory and experiment. Part 2 contains comments on how many configuration and flow parameters influence transition. Part 3 discusses some additional general aspects of transition. Part 4 discusses problems of predicting transition and comments on three prediction methods. Part 5 contains some general guidelines for prediction methodology. Keywords: Boundary layer transition, Boundary layer stability, Hypersonic boundary layers.

An analytical approach toward numerical calculation of the three-dimensional turbulent boundary layer on a sharp cone at incidence under supersonic and hypersonic flow conditions is presented. The theoretical model is based on implicit finite-difference integration of the governing three-dimensional turbulent boundary-layer equations in conjunction with a three-dimensional scalar eddy-viscosity model of turbulence. Comparison is made of present theory with detailed experimental measurements of the three-dimensional turbulent boundary-layer structure (velocity and temperature profiles), the surface streamline direction (obtained via an oil-flow technique) and surface heat-transfer rate.