Investigation of Shock Wave - Boundary Layer Interactions in a 3.57 Aspect Ratio Wind Tunnel

Abstract

A shock wave boundary layer interaction occurs where a shock wave meets any part of a surface where a boundary layer is present. A common application is the inlet of a supersonic jet engine, which can experience stagnation pressure loss, distorted flow at the engine face, thickening of the boundary layers, and in some cases, flow separations with local flow reversals. Previous experimental studies addressing shock wave boundary layer interactions (SWBLIs) have suffered from centerline separation and corner flow effects, which is the region where the boundary layers of two walls meet. The present study presents the design and validation of a new experimental test facility intended to produce centerline flows with minimal influences of the separations caused by right-angle corners in supersonic flow. SWBLIs are captured and quantified using surface static pressure measurements and a Schlieren optical flow visualization system. An accompanying computational fluid dynamics (CFD) study is also underway to provide further understanding of the interactions.