Compressor cascade flow with strong shock-wave/boundary-layer interaction

Abstract

A two-dimensional multiblock Navier-Stokes solver has been successfully applied to a supersonic compressor cascade flow with strong shock-wave/boundary-layer interaction. The cascade model investigated is typical of a fan blade section with precompression design, low flow turning, and a static pressure ratio around 2.0. Numerical calculations have been performed on a very fine grid to guarantee sufficient resolution of the supersonic/transonic flowfield with a complex wave pattern and shock-induced boundary-layer separation. High accuracy and robustness of the numerical scheme is achieved by the second-order Roe upwind-total variation diminishing scheme and a one-equation turbulence model. Investigations were performed for inlet Mach numbers from 1.28 to 1.53 and a Reynolds number of 2.6 x 106. The results are compared with those of previous experiments in a supersonic cascade wind tunnel that provide profile Mach number distributions, wake traverse data, and measured suction surface boundary-layer profiles throughout the strong interaction region. Furthermore, calculated discontinuities of Mach number and flow direction across the shock system at blade passage entrance are compared with laser anemometer data