Experimental and Numerical Investigations of Shock-Wave Boundary Layer Interactions in a Highly Loaded Transonic Compressor Cascade

Experimental and numerical investigations were conducted to investigate the variations of shock-wave boundary layer interaction (SBLI) phenomena in a highly loaded transonic compressor cascade with Mach numbers. The schlieren technique was used to observe the shock structure in the cascade and the pressure tap method to measure the pressure distribution on the blade surface. The unsteady pressure distribution on blade surface was measured with the fast-response pressure-sensitive paint (PSP) technique to obtain the unsteady pressure distribution on the whole blade surface and to capture the shock oscillation characteristics caused by SBLI. In addition, the Reynolds Averaged Navier Stokes simulations were used to compute the three-dimensional steady flow field in the transonic cascade. It was found that the shock wave patterns and behaviors are affected evidently with the increase in incoming Mach number at the design flow angle, especially with the presence of the separation bubble caused by SBLI. The time-averaged pressure distribution on the blade surface measured by PSP technique showed a symmetric pressure filed at Mach numbers of 0.85, while the pressure field on the blade surface was an asymmetric one at Mach numbers of 0.90 and 0.95. The oscillation of the shock wave was closely with the flow separation bubble on the blade surface and could transverse over nearly one interval of the pressure taps. The oscillation of the shock wave may smear the pressure jump phenomenon measured by the pressure taps.