Effect of radial inlet distortion on aerodynamic stability in a high load axial flow compressor

Abstract

Radial inlet distortion induced by boundary layer separation can significantly affect the aerodynamic performance of the compressor. The effect of radial inlet distortion on the flow structure is numerically investigated in a high load axial flow compressor. The inlet boundary is determined by the experimental results at the downstream of the distortion generator. The results reveal that tip distortion leads to a reduction in the stall margin, whereas hub distortion extends it. Radial distortion redistributes the main flow toward undistorted region due to the obstructive effect of lattice ring. The deficit in axial velocity with tip radial distortion causes the rotor to operate at a higher incidence angle near the casing, while hub radial distortion alleviates the tip blade loading. The detailed three-dimensional flow field analysis indicates that increased blade loading with tip distortion shifts the trajectories of both primary and secondary tip leakage flow toward the leading edge, thereby expanding the blockage region. Conversely, hub radial distortion unloads the rotor tip region, thereby reducing the blockage region induced by tip leakage flow. Additionally, with hub distortion, the location of separation line on the blade suction surface moves closer to the leading edge, and the flow separation around the trailing edge is intensified.