In a boundary layer ingestion (BLI) propulsion system, the compressor rotor operates under the condition of a severe inflow distortion because the boundary layers on the surface of aircraft have been ingested. In this situation, the low-momentum fluid would inevitably raise the blade load and aggravate the local aerodynamic losses, thus severely deteriorating the compressor overall aerodynamic performance. The paper has applied a sweep optimization design based on surrogate model and genetic algorithm to a transonic axial-flow compressor at the condition of total pressure distorted inflow. The optimization objective is to improve the compressor efficiency. The sweep design has been implemented through controlling the shape of the blade stacking line and the time-saving steady numerical simulations are mainly utilized in the optimization process. Then, the full-annulus unsteady CFD calculations have been conducted for validating the optimized design and further analyzing the flow mechanisms of performance improvements. The computational results show that the inflow distortion has led to non-uniform distributions of the tip leakage loss and the hub flow separation loss along the circumferential direction. It is found that the rotor blade has suffered from the highest tip leakage losses and hub separation losses during the blade leaving the distorted region. After rotor sweep optimization, both tip leakage and hub separation losses have been notably reduced along the full circumferential direction. Moreover, for different pitchwise locations, the optimized rotor sweep design has achieved more aerodynamic performance improvements in the process of the rotor blade leaving the distorted region than those in the processes of the rotor blade moving out of, moving towards and entering the distorted region. Overall, the optimized design performs better than the baseline design in terms of adiabatic efficiency in the entire operating range. In the meanwhile, the total pressure ratio has been maintained. Specifically, an approximately 0.5% increment of efficiency has been achieved by the optimized design at the optimization point. For this paper, the major purpose is to offer useful guidelines for fan/compressor design strategy to achieve high aerodynamic performances under distorted inflow working condition.

在边界层进气（BLI）推进系统中，由于飞机表面上的边界层已被摄入，因此压缩机转子在严重的进气扭曲情况下运行。在这种情况下，低动量的流体将不可避免地增加叶片载荷并加剧局部空气动力损失，从而严重降低压缩机的整体空气动力性能。在总压变形入流条件下，将基于替代模型和遗传算法的扫掠优化设计应用于跨音速轴流压气机。优化目标是提高压缩机效率。通过控制叶片堆叠线的形状来实现扫掠设计，并且在优化过程中主要利用了省时的稳态数值模拟。然后，进行了全环面非稳态CFD计算，以验证优化设计并进一步分析性能改进的流程机制。计算结果表明，流入畸变导致尖端泄漏损失和轮毂流分离损失沿圆周方向的不均匀分布。已经发现，在叶片离开变形区域期间，转子叶片遭受最大的尖端泄漏损失和轮毂分离损失。经过优化的转子扫掠后，叶尖泄漏和轮毂分离损失在整个圆周方向上均得到显着降低。此外，对于不同的俯仰方向位置，与转子叶片移出，移向并进入变形区域的过程相比，优化的转子扫掠设计在转子叶片离开变形区域的过程中实现了更多的空气动力学性能改进。总体而言，就整个操作范围内的绝热效率而言，优化设计的性能要优于基线设计。同时，总压力比得以保持。具体而言，通过优化设计在优化点获得了约0.5％的效率增量。对于本文，主要目的是为风机/压缩机的设计策略提供有用的指导，以在变形的流入工况下实现较高的空气动力学性能。进入并进入变形区域。总体而言，就整个操作范围内的绝热效率而言，优化设计的性能要优于基线设计。同时，总压力比得以保持。具体而言，通过优化设计在优化点获得了约0.5％的效率增量。对于本文，主要目的是为风机/压缩机的设计策略提供有用的指导，以在变形的流入工况下实现较高的空气动力学性能。进入并进入变形区域。总体而言，就整个操作范围内的绝热效率而言，优化设计的性能要优于基线设计。同时，总压力比得以保持。具体而言，通过优化设计在优化点获得了约0.5％的效率增量。对于本文，主要目的是为风机/压缩机的设计策略提供有用的指导，以在变形的流入工况下实现较高的空气动力学性能。通过在优化点进行优化设计，效率提高了5％。对于本文，主要目的是为风机/压缩机的设计策略提供有用的指导，以在变形的流入工况下实现较高的空气动力学性能。通过在优化点进行优化设计，效率提高了5％。对于本文，主要目的是为风机/压缩机的设计策略提供有用的指导，以在变形的流入工况下实现较高的空气动力学性能。