In this paper, large eddy simulation (LES) was performed to investigate the shock wave/boundary layer interaction (SBLI) phenomenon in transonic compressor cascades with a chord Reynolds number of 2.12 × 106. A comprehensive analysis was conducted on both the SBLI structures inherent to the transonic compressor cascade and the coherent vortex structures within the boundary layer. The underlying mechanisms of the shock-induced boundary layer transition and the shock low-frequency unsteadiness in the transonic compressor cascade were elucidated through spectral and dynamic mode decomposition (DMD) analysis. The results revealed that boundary layer separation induced by the SBLI cannot reattach, leading to the formation of large-scale coherent vortex structures. Spectral analysis revealed that the shock-induced boundary layer transition in the transonic compressor cascade was dominated by inviscid Kelvin–Helmholtz (K–H) and secondary instability mechanisms, characterized by a dimensionless Strouhal number of 0.06. Additionally, pressure signals showed the variations in sub-frequency from the separated shear layer to the main flow. The oscillation amplitude of the shock foot was significantly greater than that of the shock main body, and the oscillation frequency of the shock foot was consistent with the sub-frequency. The oscillation frequency of the shock main body coincided with that of the compression ramp and flat plate configurations. Finally, DMD modal analysis indicated that high-frequency modes were correlated with turbulent fluctuations in the boundary layer, while medium- and low-frequency modes corresponded to shedding motion in the separated shear layer and low-frequency motion of the shock. This work promotes the understanding of the complex flow mechanisms of SBLI in the transonic compressor cascades.

In this paper, a two-dimensional transonic cascade of German Aerospace Center (DLR) was redesigned with multi-objective optimization. The aerodynamic performance of the cascade was optimized with inlet Mach number of 1.1 by using the blade design method with thickness distribution superposed on the camber line. The non-dominated sorting genetic algorithm, (NSGA) II multi-objective optimization algorithm, was used to determine higher performance cascades. Cascade optimization was carried out within a unique inlet Mach number by varying the shapes of both the camber line and the thickness distribution curve. To enable the cascade to operate under specific conditions, we used an inlet Mach number control loop (IMCL) to regulate the outlet boundary conditions. Optimization was conducted to minimize the cascade total pressure loss coefficient under the three inlet conditions. The optimization results indicated that the total pressure loss coefficient can be reduced by 22% under the design condition, and the range of working conditions was widened.

The effects of back pressure on the transonic cascade operating state are crucial and can determine the structure of internal shock waves and secondary flows. In this paper, numerical methods validated by experiments were employed to investigate the evolution mechanisms of the inlet flow field, shock structure, secondary flow structure, and cascade performance under different back pressures. Analysis revealed that transonic cascade exhibited unique incidence characteristics in the inlet flow field under both subsonic and supersonic regimes, although these two regimes involved different physical mechanisms. The results revealed that the operating state of the transonic compressor cascade under the unique incidence condition was influenced by the outlet back pressure, and there existed a critical static pressure ratio. The critical static pressure ratio shifted from 1.61 for two-dimensional flow to 1.37 for three-dimensional (3D) flow at M1 = 1.1, due to the corner separation and the characteristics of 3D shocks. The 3D shock structure exhibited a non-uniform distribution along the spanwise direction due to the influence of back pressure and the separated boundary layer. The vortex structures analysis revealed that the secondary flow structure on the sidewalls of the transonic compressor cascade was primarily dominated by corner vortices, whose formation mechanism was related to the interaction between the shock wave and the sidewall boundary layer. Additionally, this interaction also led to the formation of detached shock and lip shock vortex structures. Finally, loss analysis indicated that the wake region of the transonic cascade primarily includes six types of loss, and the total loss of the cascade decreased with the rise in back pressure.

In this paper, a transonic compressor cascade was optimized to improve its aerodynamic performance. A new blade parameterization method with 16 control variables was first proposed to fit the shapes of the suction and the pressure side, as well as the leading edge. Then, the Kriging surrogate-model-based genetic algorithm (GA) was used to optimize the performance of the transonic cascade. The optimization algorithm is effective in reducing the total pressure loss while extending the working range of the cascade. The results show that the total pressure loss coefficient could be reduced by 11% at the best airflow angle and the working range could be extended by 6.9% for the optimized cascade in two-dimensional simulations. Similar improvement results could also be obtained in the simulations of their linear cascade cases. Detailed analyses show that the relative maximum thickness positions of the optimized blades move forward by about 10% to the leading edge, and the radii of curvature of the front half of the suction and pressure surfaces increase, compared with the initial blade. This makes the front half of the optimized blades look more closely like a wedge. Consequently, the passage shock strength is reduced and the shock changes from the passage normal shock to oblique shock. The weakened shock strength leads to the disappearance of the flow separation caused by the shock boundary layer interaction on the suction surfaces of the optimized blades, and results in a narrowed wake width at the outlet section.

Highly three–dimensional and complex flow structures are closely related to the aerodynamic losses occurring in the transonic axial–flow compressor. The large eddy simulation (LES) approach was adopted to study the aerodynamic performance of the NASA rotor 37 for the cases at the design, the near stall (NS), and the near choke (NC) flow rate. The internal flow vortex topology was analyzed by the Q–criterion method, the omega (Ω) vortex identification method, and the Liutex identification method. It was observed that the Q–criterion method was vulnerable to being influenced by the flow with high–shear deformation rate, especially near the end–wall regions. The Ω method was adopted to recognize the three–dimensional vortex structure with a higher precision than that of the Q–criterion method. Meanwhile, the Liutex vortex identification method showed a good performance in vortex identification, and the corresponding contribution of Liutex components in the vortex topology was analyzed. The results show that the high–vortex fields around the separation line and reattachment line had high vortex components in the x–axis, the tip clearance vortices presented a high–vortex component in the y–axis, and the suction side corner vortex possessed high–vortex components in the y– and z–axes.