Experimental investigation on the performance of compressor cascade using blended-blade-end-wall contouring technology

Three-dimensional flow separations commonly occur in the corner region formed by the blade suction surface and end wall in compressors. How to control or reduce these separations is a vital problem for aerodynamic designers all the time. Blended blade and end wall contouring technology has been proposed to control flow separation for several years and validated in many cases using the numerical method, but experimental data was not obtained so far. So in this paper, the baseline cascade scaling from the NACA65 airfoil with 42 degrees turning angle is designed, tested, and analyzed firstly. Then, based on the experimental results of the baseline cascade, blended blade and end wall contouring is applied to the suction surface and hub corner region of the baseline cascade and the detailed experiment is carried out. The results show that the blended blade and end wall contouring technology can decrease the total pressure loss by 8% and 7% at 0 degrees and +10 degrees incidence angles separately. The improved span range mainly focuses on the 10-25% span height. The rolling change of the passage vortex influenced by the accumulation of low energy fluid driven by cross flow in the hub corner should be the main reason for the performance improvement.