Experimental investigation of two-dimensional shock boundary layer interactions

Experiments were performed on the interaction of oblique shock waves with flat plate boundary layers in the forward test section of the 30.48 cm. x 30.48 cm. (1 ft. x 1 ft.) Supersonic Wind Tunnel at the NASA-Lewis Research Center. Measurements of the plate surface static pressure and shear stress distributions as well as boundary layer velocity profiles were obtained through the interaction region. Flow conditions for these investigations encompassed a Mach number set from 2.0 to 4.0 and an overall unit Reynolds number range of 4.72E6 to 2.95E7 /m. The findings presented herein provide a complete description of two-dimensional interactions with initially laminar boundary layers over the Mach number range 2.0 to 4.0. Additional information with regard to interactions involving initially transitional boundary layers is provided over the Mach number range 2.0 to 3.0 and those for initially turbulent boundary layers at Mach 2.0. The experiments were motivated with achievement of benchmark results as their primary goal, in order to yield detailed information of high accuracy for service as a test case of analytic and numeric schemes of the present and future.