民用航空发动机过渡段的流动情况对压缩系统整体性能有着重要影响。探究过渡段内部流场的流动机理和损失机制有利于高性能压气机的研制。本文使用数值模拟的方法对某航空发动机增压级过渡段进行了流场分析，并与实验结果进行对比验证。结果表明:过渡段出口流场在20%-90%径向位置处分布较为均匀，气流的总压损失主要集中在支板尾缘和上下壁面附近，主要损失机理包括下壁面曲率引起的附面层分离和支板表面附面层分离，角区的二次流扰动沿流向始于下壁面曲率最大处，总压损失会随流速的增加而增大，流速越大的总压损失增量也会越大。

The flow situation in the compressor intermediate casing duct of an aircraft engine compressor has a sig-nificant impact on the overall performance of the compressor. Exploring the flow mechanism and loss mechanism of the internal flow field in the intermediate casing duct is beneficial for providing a solid foundation for the development of high-performance compressors. This article conducts numerical simula-tion and flow field analysis on the model obtained through geometric parameterization of the intermediate casing duct profile, and compares it with experimental results. The results show that the flow field at the exit of intermediate casing duct is evenly distributed in the radial position of the 20% -90% flow channel, and the total pressure loss at the exit is mainly concentrated near the trailing edge of the support plate and the upper and lower walls. The loss in the flow channel is caused by the combined curvature of the lower wall and the boundary layer on the support plate surface, The secondary flow disturbance in the corner re-gion starts along the flow direction at the point where the curvature of the lower wall is maximum, and the total pressure loss will increase with the increase of flow velocity. The larger the flow velocity, the greater the increment of total pressure loss.

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