为了解决压气机级间泄漏与二次流流动问题，航空发动机轴流压气机静叶根部与转子之间通常采用篦齿进行封严。为研究封严篦齿泄漏流对压气机性能的影响，基于某轴流压气机建立了带封严篦齿真实结构的几何模型，采用三维数值模拟的方法，研究了篦齿泄漏流对某轴流压气机主流涡系结构和流动损失的影响，并探究了其影响机理。结果表明：封严篦齿泄漏流使压气机的压比和效率都有不同程度的下降；篦齿泄漏会增强上游转子叶根吸力面的尾缘角区涡和静子叶根吸力面的马蹄涡，并使设计工况的上游转子和静子的流动损失分别增大3.1%和13.1%；静子叶根后附面层低能流体被抽吸，改善了下游流场，使下游转子流动损失减小2.4%；在近喘振点，由于压气机内流场恶化严重，篦齿泄漏带来的流场变化并不显著，泄漏流对主流影响小。

In order to solve the problems of leakage between stages and secondary flow in the compressor, labyrinth seals are usually used between the root of stator vane and the rotor of an aero engine axial compressor. To investigate the impacts of leakage flow of labyrinth seals on performance of compressor, a geometric model of the real structure of seals is established based on an axial compressor, and the effect of the leakage flow of seals on the flow loss of the main vortex of an axial compressor and its influence mechanism are studied by using the method of threedimensional numerical simulation. Results show that the pressure ratio and efficiency of the compressor reduce to different extents, which are attributed to leakage flow of labyrinth seal. Besides, it enhances the trailing edge separation vortex at the root of suction side of the upstream rotor blade and the horseshoe vortex at the root of suction side of the stator vane. At design point, the flow loss of the upstream rotor blade and the stator vane are increased by 3.1% and 13.1% respectively; meanwhile, the low energy fluid in the boundary layer at the outlet of the stator vane′s root is pumped to improve the downstream flow field. The flow loss of downstream rotor blade is reduced by 2.4%. In addition, due to the serious deterioration of flow field in the compressor near the surge point, labyrinth seal leakage leads to insignificant flow field changes and the leakage flow has a negligible impact on the mainstream.

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