HYPERSONIC FLOW OVER BLUNT SLAB WINGS WITH SWEEP ANGLES FROM 0 TO 70 DEGREES.

An experimental and theoretical investigation was made to determine the pressure distributions over swept wings in hypersonic flow. Two models having hemicylindrical leading edges were tested in the Aerospace Research Laboratories 20-inch hypersonic wind tunnel. Tests were made over an angle-of-attack range of 25 degrees at five-degree increments for the sweep angles 0, 50, 60, and 70 degrees at a nominal test Mach number of 14 and a Reynolds number of 35,000 based on the nose diameter of the models. The test data measured on the nose region was in fair agreement with the theoretical pressure distributions for each sweep angle and appeared to be independent of Mach number as expected. At angles of attack up to 15 and 25 degrees, the flow over the expansion and compression surfaces appeared to be two-dimensional. Fairly good agreement was obtained between the measured test data and an inviscid theoretical analysis based on a modification of the blast wave theory. Author