Experimental study of wall temperature effect on shock wave/turbulent boundary layer interaction in hypersonic aircraft

The wall temperature effects on turbulent boundary layer and shock wave/turbulent boundary layer interaction are studied experimentally at Ma 2.7 with wall-to-recovery temperature ratio Tw/Tr from 0.67 to 1.2. The high-resolution turbulent flow structures are captured and analyzed by applying the nano-tracer planar laser scattering and particle image velocimetry techniques. The influence mechanisms of wall temperature on the vortices and interaction flow field are explained. The results indicate that under the heating wall condition, the increase of gas viscosity and volume causes the turbulent boundary layer to develop thicker and faster. The thickness increases 0.79 mm as Tw/Tr increases 0.2. The boundary layer shape presents intermittence with large-scale vortices increasing in size and number. The shock wave interaction point moves 1.43 mm upstream as Tw/Tr increases 0.2 due to the increasing boundary layer thickness. More intensive separation of the boundary layer leads to the height and length extension of the separation region and the increase of the separation shock angle. Under the wall cooling effect, the boundary layer thickness decreases, and the global vorticity reduces. The velocity field becomes more uniform. The present study can provide an important experimental reference for the design of hypersonic aircraft.