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.

In a boundary layer ingesting (BLI) propulsion system, the fan is continuously exposed to inflow distortions. The distorted inflows lead to nonuniform loss distributions along the radial and circumferential directions. Since the rotor tip suffers from higher intensive distortion, the local loss increment is a major contributor to the BLI fan performance penalty. To explore the effects of distorted inflows on tip leakage flow evolutions and associated mechanisms for increased loss in a BLI fan, three-dimensional full-annulus unsteady simulations are conducted. Results show that about 54% of total additional losses due to distortion are formed in the tip region and more than 80% of tip region entropy generation is related to the tip leakage flow. The intensities of leakage vortex–shock interactions vary at different annulus locations. When the rotor moves into the distorted region, the vortex–shock interaction is weaker than the undistorted locations due to attenuated leakage flow. At the locations where the rotor is moving out from the distorted region, the vortex–shock interaction is notably enhanced because the front part of the blade tip airfoil suffers a higher load, resulting in a rapid vortex core expansion and eventually vortex breakdown. The increase of flow blockage in the front section of blade tip passages at local circumferential positions leads to a corresponding rise of flow loss. The findings in this study highlight the impacts of tip leakage flow on aerodynamic loss of fan working under BLI inflow distortion and provide improved understandings of loss mechanisms in a BLI fan.

Axial-flow compressors work against varying inlet boundary layers in real working conditions and are therefore required to perform well and robustly. This paper presents a surrogate-based optimization procedure applied to a transonic compressor to improve its efficiency and reduce the sensitivity of efficiency variation to uncertain inlet boundary layer thicknesses while maintaining the total pressure ratio. The aerodynamic optimization of compressors involves high-fidelity computational models that would cost high amounts of computational time. To implement the optimization, a region-segmentation combining surrogate model is used that is based on combinational use of the region-segmentation idea and combining surrogate modeling method to further improve prediction accuracy and reduce computational cost. Based on the region-segmentation combining surrogate model, an optimization procedure is constructed and applied to a transonic compressor. The computational results of the benchmark function and compressor optimization indicate the validity of the region-segmentation combining surrogate model in improving the prediction accuracy and computational efficiency. The optimization procedure also presents the ability to improve the compressor efficiency and make the compressor perform well and robustly at uncertain inlet boundary layer thicknesses while maintaining the total pressure ratio. The achieved aerodynamic benefits of the compressor have demonstrated the feasibility and effectiveness of the optimization strategy.