Experimental study of a mach 3 bump compression flowfield

The flowfield surrounding a bump compression surface in supersonic flow has been studied experimentally. The compression surface was mounted in a blowdown-type wind tunnel that was operated at a Mach number of 2.95 with a unit Reynolds number of 39 x 10 m^-1. Qualitative visualization of the flowfield included the use of oil streakline patterns to capture the surface flow trends and schlieren photography to view shock waves and boundary layer evolution along the bump surface. Static pressure measurements were obtained along the bump's centerline and at four spanwise planes. Along these same planes of investigation, laser Doppler velocimetry was used to provide mean velocity and turbulent stress distributions. The experiments showed that the bump induces a curved shock system originating from the leading edge of the compression surface without causing flow separation. The well-faired contour of the bump counteracts the initial pressure rise caused by the shock, yielding favorable pressure gradients in both the streamwise and spanwiseoutward directions. These gradients push low-momentum boundary layer fluid away from the model centerline to the sides of the bump. The net result of this is a downstream boundary layer along the bump centerline that is thinner (by a factor of two) and less turbulent than the incoming boundary layer.