Axial flow compressors operate under adverse pressure gradient condition. The growth of boundary layers on blade susfaces and end walls causes increased total pressure loss and flow blockage. This paper presents a method to reduce the pressure loss through suction of low momentum end wall boundary layer in a high Mach number, high turning compressor stator cascade. The end wall suction was applied though slots at discrete chordwise locations near the blade suction surface. CFD simulations were carried out for different slot locations, different suction mass flow rates and at different incidence angles using ANSYS FLUENT software at design inlet Mach number of 0.762. The boundary layer suction technique significantly reduced the total pressure loss across the cascade by reducing the separated flow region at the end wall. The end wall suction slot located at 50–75% blade chord was found to be most effective in reducing the total pressure loss coefficient by 54.6% compared to the baseline case without end wall suction.
Transonic Axial Compressor, Boundary Layer Suction, Controlled Diffusion Aerofoil, Total Pressure Loss, Flow separation
