An experimental investigation of the forebody of a hypersonic inlet model and a comparison with theory

Abstract

An experimental investigation of the flow field of a hypersonic inlet forebody model has been conducted in a Mach 4 airstream. Test section Reynolds number was approximately 6.6 X 107 per meter at a stagnation pressure of 13.6 atmospheres. The model consisted of two 8° wedges in the same plane and sloping toward each other, whose leading edges made angles of 55° to the approaching flow. The side walls were designed to follow the flow streamlines on the wedge surfaces and the shock waves from the wedge leading edges. The data was obtained from surface static pressure taps, pitot probe surveys, oil-streak tests, and schlieren photographs. The results of the investigation indicate that the leading-edge shock waves join in a manner producing a region of three-dimensional, conical-like flow near the model's plane of symmetry. The flow and shock wave structure in this region were similar to the corner flow situation and the region size was underpredicted by the theoretical analysis. In the analysis the free-stream flow was broken into velocity components normal and tangential to the leading edge and utilized two-dimensional methods. It was possible to predict the flow properties and shock wave structure with confidence outside the center region, although the experimental Mach number was approximately 4 percent less than the calculated value due to viscous effects on the model surface. An appendix is included which uses the analysis to make a parametric study of several swept wedge configurations.