Unsteady flow phenomena are common in turbomachinery and have serious impacts on its stable operation. This paper reports a self-induced unsteady phenomenon inside a centrifugal impeller due to strong secondary flow that has not been previously reported. A high-loading centrifugal impeller is selected as the research object, an accurate three-dimensional model is established, and exhaustive numerical simulations are performed based on unsteady compressible Navier–Stokes equations. The unsteady performance of the compressor is measured in detail during the experiment, and the experimental results are compared with the numerical simulation results to verify the reliability of the numerical simulation method. The results of the unsteady pressure signal show the existence of unsteady phenomenon when the operating point is far away from the choke boundary. The power spectrum results of the unsteady pressure distribution at the impeller blade tip are analyzed, and two different modes of this unsteady phenomenon are confirmed. Higher-order frequency characteristics are not observed near the stall boundary. The detailed analysis of the flow field inside the impeller clarifies that the instability of the interface formed between the secondary flow and the main flow induces this unsteady phenomenon. This unsteady pressure disturbance propagates inside the blade passage through pressure waves, which contributes to the instability of the interface and the existence of this self-induced unsteady phenomenon.

High-pressure ratio centrifugal compressors’ diffusers face challenges from high-velocity, non-uniform flow at the impeller outlet, decreasing efficiency and stall margin. To address this, this paper presents a novel vaned diffuser passage design method that successfully improved the compressor’s performance. An optimization method using axisymmetric hub contours and NURBS curves was applied to modify the diffuser’s design. After optimization, centrifugal compressor peak efficiency increased by 0.78%, and stall margin expanded from 12.8% to 20.4%. Analysis at the peak efficiency point showed loss reduction mainly from decreased recirculation and mixing losses in the diffuser’s vaneless and semi-vaneless spaces. Furthermore, correlation analysis and Mach number distribution revealed that flow behavior at the diffuser’s leading edge significantly influences efficiency. Consequently, design principles emphasize satisfying specific Mach number distribution rules at the diffuser’s leading edge under certain inflow conditions for optimal performance.

To explore the internal flow mechanism and improve the performance of a supersonic compressor, an efficient global optimization design method was developed for an axial flow compressor and applied in the optimization design of a prototype supersonic compressor. Based on the multiple circular arc (MCA) blade parameters, the method can be used to parameterize the elementary stage of the blade. The optimized solution is obtained by changing the elementary stage and stacking lines of the blade during the optimization process. It has the advantages of fewer optimization variables, strong physical intuition, and a smooth surface. The optimization results show that a change in the rotor blade shape parameters has a significant effect on the compressor efficiency under design conditions, while a change in the skewed-swept parameters of the stator is the main factor that improves the compressor’s performance under near-stall conditions. Further numerical results show that the optimized rotor changes the form of the shock, weakens the degree of shock boundary layer interference, inhibits the radial migration flow of the supersonic rotor, reduces the loss of the rotor blade top, and improves the performance of the compressor under design conditions. The stator’s optimization restrains the generation of a concentrated shedding vortex at the root of the blades and greatly improves the stall margin of the compressor. Finally, the total pressure ratio and flow rate are less than 1% of the values based on the prototype operating conditions, the design mass flow of the optimized high-load supersonic compressor is increased by 0.25%, the isentropic efficiency is increased by 1.05%, and the stall margin is enhanced by 3.5%, thus verifying the effectiveness of the optimization method.

To comprehensively explore flow control method of profiled end wall for turbine stage at low Reynolds numbers, a surrogate model optimization platform including non-uniform rational B-spline surface parameterization method, support vector regression, and improved chaos particle swarm optimization algorithm is integrated. Optimization designs have been carried out for stator profiled end walls, rotor profiled end wall, and combined end walls, respectively. The results indicate that under the constraint of the output power, the application of various profiled end wall design cases all can effectively improve the aerodynamic performance of the turbine stage. By organizing the flow field of downstream rotor, the profiled end wall of stator can significantly affect the stage efficiency. The flow control benefits of the profiled end wall of the rotor is from the obstruction of the cross migration of the pressure side leg of the horseshoe vortex. The application of profiled end wall on stator has the most practical engineering value. Self-organizing maps and Shapley methods are used to explore potential correlation information of aerodynamic parameters and summarize design experience. The sensitive design variables of profiled end walls are extracted. Based on the local controllability of NURBS surfaces, the regions that affect the stage efficiency are mainly concentrated in the middle of the stator passage, near the stator trailing edge and near the rotor leading edge. The regions with a significant impact on the output power of the turbine stage are near the trailing edge of the rotor and stator. The corresponding design rules of end walls modeling are summarized.