This paper presents a numerical study on three-dimensional flow phenomena near the endwall of a linear high-turning compressor cascade at supercritical flow conditions. The compressor cascade with 60° camber angle was designed at a higher supercritical speed (M1>0.9) by optimum method based on the baseline which aimed at improving the flow near the stator hub of small transonic fans. The camber line and thickness distribution curves of the baseline are formed by quadratic polynomials and double cubic curves respectively. The stack line and the thickness distribution near the end-wall were chosen as optimization variables to approach the objective function of total pressure loss coefficient, since they are the two main geometry parameters which can influence end-wall flow obviously. The analysis in current paper focuses on comparing the flow phenomena near the end-wall of baseline cascade with that of optimized one. Numerical simulation results are presented to show the loss reduction from the baseline to the optimized cascade near end-wall. The boundary-layer development on the suction surface, flow separation structure, shockwave and local supersonic area on the suction surface near the end-wall are analyzed in detail. The optimized cascade has a stronger shockwave near the leading edge. It was found that the radial flow of the boundary-layer caused by the optimization of stack line is the key factor influencing the aerodynamics loss near the end-wall at supercritical condition which also plays an important part in second-flow and flow separation in the corner. An understanding of the low-loss pattern of the end-wall flow and the flow filed structure for high-turning compressor at higher supercritical flow conditions then is summarized at the end of this paper.

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