Investigation of a turbulent boundary layer on a hypersonic aircraft model

An algorithm for calculation of a spatial compressible turbulent boundary layer on the surface of a pointed body is developed. The algorithm is based on the numerical solution of three-dimensional equations and algebraic models of turbulence. The flow around a hypersonic aircraft model is calculated, and the resultant Stanton numbers are compared with experimental data. The influence of the Mach number, the angle of attack, and the Reynolds number on the boundary-layer parameters is studied. It is shown that the change in the location of the transition zone has a weak effect on the skin-friction coefficient in the region of developed turbulent flow.