Investigations of shock wave boundary layer interaction by panoramic methods

Interaction of a shock wave with the boundary layer on a half-airfoil model is studied. The experiments are performed in a wind tunnel with the free-stream Mach number M $\approx$ 0.75 and stagnation pressure $P_0$ = 10$^5$ Pa. The half-airfoil model is mounted on the wall of the test section of the wind tunnel. Pressure distributions over the model surface are obtained by a method with luminescent pressure transducers and by a method with pressure taps. The limiting streamlines on the model are visualized, and thermographic visualization is also performed. For experimental parameters, numerical simulations are performed with an approach based on using the Reynolds-averaged Navier–Stokes equations. The three-dimensional structure of the flow is analyzed, and significant differences in the measured and simulated results for the flow in corner separation regions are revealed.