Abstract
The interaction of a shock wave with a boundary layer on a half-airfoil model has been studied. Experiments were performed in a wind tunnel at a freestream Mach number \( \approx {\kern 1pt} 0.75\) and a stagnation pressure \({{P}_{0}}{{ = 10}^{5}}\) Pa. The half-airfoil model was mounted on the wall of the test section of the wind tunnel. Pressure distributions on the model surface were obtained using pressure sensitive paints and pressure taps. Visualization of the surface streamlines on the model by means of oil flow visualization technique and measurements of surface temperature distribution using infra red CCD camera were performed. Numerical simulation of the flow was performed with experimental parameters using the Reynolds-averaged Navier–Stokes approach. The three-dimensional flow structure was analyzed, and significant differences were found between the measurement results and numerical simulation data for flow in the corner separation regions.
