3-D transonic flow in a compressor cascade with shock-induced corner stall

An experimental and numerical study of the transonic flow through a linear compressor cascade with endwalls was conducted. The cascade with or low aspect ratio of 1.34 was tested at an inlet Mach number of 1.09 and a Reynolds number of 1.9 x 10(6). Detailed flow visualizations on the surfaces and five-hole probe measurements inside the blading and in the wake region showed clearly a three-dimensional boundary layer separation on the blade surface and the sidewall, and a severe corner stall induced by a strong 3-D shock system tit blade passage entrance. The experimental data have been used to validate and improve the 3-D Navier-Stokes code TRACE. Results showed an excellent resolution of the complex flow field. Surface pressure distribution, on the entire blade surface and the endwalls, flow angle and total pressure contours within the blade passage and the wake are compared with the experimental results, Air analysis, of the secondary flow of this highly staggered cascade did not show the classical corner vortex. Instead, a severe flow deviation and partly reverse flow, near the walls is wen. The flow solver helped to identify a weak ring vortex that originates from the passage sidewall. Surface oil flow pictures on the blade contour and the, sidewall are tit qualitatively good agreement to numerical surface streaklines, A considerable improvement of the numerical results could he achieved by a gradual grid refinement, especially in the corner region and by successive code development.