Experimental study of aerodynamic heating in the region of an incident shock wave boundary layer interaction

The effect of an incident shock wave on the thermal and dynamic parameters of a supersonic flow on a flat plate is experimentally studied. The shock wave was initiated by a wedge with an opening angle of 12 degrees. The oncoming flow Mach number was 2.48, the total pressure was 530 kPa, the total temperature was 295 K. The Reynolds number based on the boundary layer length from the nozzle throat to the beginning of the test section was at least 2 center dot 107. The static pressure distributions along the flat plate length are presented for two flow patterns with a shock wave and without any disturbances. The 2D2C-PIV method was used to obtain the velocity field in the wind tunnel test section and in the boundary layer separation region on the flat plate. Applying PIV method allowed to register the main characteristics of the interaction region including mean velocity field, incident and reflected shock wave patterns, as well as the boundary layer distortion. The cooling rates of the flat plate were obtained using an infrared camera, which made it possible to calculate the distributions of the temperature recovery factor and the heat transfer coefficient along the flat plate surface. An increase in the temperature recovery factor was noted for flow with an induced shock wave in comparison with an undisturbed flow. At the same time, an uneven distribution of temperature recovery factor and heat transfer coefficient was observed in the separation region. The results of the study can be useful in calculating high-speed channel flows with heat exchange in air- and spacecrafts propulsion systems.