High-pressure ratio is one of the important characteristics of the sustainable development of the modern aero-engine compressor components. When the fluid flows through the compressor cascade row, it will be influenced by both the streamwise pressure gradient and the transverse pressure gradient, which will cause hub-corner separation or stall. In this paper, different diffusion factors are chosen for the cascades. Each diffusion factor has different turning angles. The formation mechanism of hub-corner separation is studied under the condition of zero angle of attack. Numerical simulation is used to study the influence of pressure gradient on the flow field in the corner. The scale of the concentrated shed vortex forms in the suction surface increases with the increasing of the transverse pressure gradient during the hub-corner separation. When the streamwise pressure gradient increases, the suction surface vortex forms the corner stall. By reasonable design, the two vortexes can cancel out each other. At this time, the loss of cascades is the minimum.

Based on the flow mechanism of the corner separation/stall, the trailing gaps are set on three typical turn angle cascades. The results show that the trailing gaps can control the radial development of the suction surface vortex during the stall and improve flow field. The jet cannot blow the suction side boundary layer away during the corner separation, because the gap does not change the static pressure distribution at the root of the cascade. In a word, the trailing edge gaps can not only inhibit the separation in the hub corner but also have the minimum leakage loss at design point. It can be used as an effective and practical compressor design method.

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