Evolution of unsteady secondary flow structures near the onset of stall in a tip-critical transonic axial flow compressor stage

The aerodynamic stability of an axial compressor stage depends on the rotor and stator. However, due to the specific design requirements, the evolving flow field and the resulting secondary flow structures are different in both components. This study investigates the evolution of dominant secondary flow structures occurring in the rotor and stator of a tip-critical transonic compressor stage at the near-stall condition using unsteady numerical analysis and validates performance characteristics with the experimental data. The investigation reveals that the presence of rotor tip shock creates a large difference in the pressure gradient across the pressure surface and the suction surface, intensifying tip leakage flow and shock-induced boundary layer separation. The higher incidence angle near the hub leading edge creates a local separation and reattachment zone. The radial pressure gradient causes the low momentum flow from this local separation zone to migrate radially upward. This migrated flow interacts with the tip leakage flow and separated blade boundary layer, eventually creating a colossal recirculation zone and subsequent rotor blockage of around 46%. The increasing streamwise adverse pressure gradient pushes the tip shock upstream, and at the onset of the stall, the flow directly separates from the rotor leading edge avoiding the shock formation. The stator flow field is dominated by the asymmetric hub corner separation induced by the streamwise adverse pressure gradient and the tip corner separation caused by the vortex structures convected downstream from the rotor tip region leading to stator blockage of around 48%. Along with the blade passing frequency, four other dominating frequencies (0.06×BPF, 0.12×BPF, 0.44×BPF, and 0.84×BPF) related to the aerodynamic instabilities are observed at the inception of stall.