Post-Shock Expansion Phenomenon Caused by Normal Shock/Turbulent Boundary Layer Interaction

Abstract

When interaction between a normal shock wave and a turbulent boundary layer occurs on the curved surfaces of transonic or supersonic airfoils, the static pressure in the vicinity of the boundary layer edge immediately downstream of the shock wave decreases along the stream lines. This phenomenon, the so-called post-shock expansion, has been studied by many researchers since it has a significant effect on the airfoil performance. Normal shock/turbulent boundary layer interaction may also be found in such confined flows as supersonic inlets or diffusers. The present study describes an experimental investigation of the post-shock expansion in supersonic diffuser flows. The Mach number immediately upstream of the shock wave is varied from 1.10 to 1.85 and the Reynolds number, based upon the boundary layer thickness of the approaching flows, is 1.3-2.5 x 10⁴. Measurement methods include schlieren optical observations, surface oil tracers, static pressures, and pitot pressures in the boundary layer. The effect of the approaching flow Mach number on the post-shock expansion phenomenon is presented.