A numerical investigation of the effect of end-wall boundary layer skew on the aerodynamic performance of a low aspect ratio, high turning compressor cascade

The paper reports on a numerical investigation into the effects of inlet boundary layer skew on the aerodynamic performance of a high turning 50 deg, 2D compressor cascade. The cascade geometry is representative of stator hub sections in highly loaded single-stage axial-flow low-speed compressors. 2D blades with NACA 65 thickness distribution on circular arc camber lines were used. The blade aspect ratio was 1.0, the space/chord ratio 0.5 and the stagger angle 25 deg. The simulations were done with a commercially available, steady three-dimensional RANS solver with the Spalart-Allmaras turbulence model. The incoming end-wall boundary layers were assumed to be collateral or skewed. In both cases the profile boundary layers were filly turbulent. The Reynolds-number was fixed at 600000 and the thickness of the incoming end-wall boundary layer was 0.1. Results are shown for an inlet-air angle of 50 deg, representing the impact free inlet-air angle of a hypothetical cascade with zero-thickness blades. Contrary to what has been expected, the results do not show (hub) corner stall, neither with nor without end-wall boundary layer skew. Flow reversal happens to occur almost exclusively on the suction surface of the blades, not on the end-walls. The end-wall flow is high v overturned, when the incoming boundary layer is collateral and is much less curved when the incoming boundary layer is skewed and (re)energized. This in turn leads to an interaction between the end-wall and blade suction surface flow which is much stronger in the first than in the second case with corresponding higher and lower losses, respectively.