Due to the complexity of the internal flow field of compressors, the aerodynamic design and optimization of a highly loaded axial compressor with high performance still have three problems, which are rich engineering design experience, high dimensions, and time-consuming calculations. To overcome these three problems, this paper takes an engineering-designed 2.5-stage highly loaded axial flow compressor as an example to introduce the design process and the adopted design philosophies. Then, this paper verifies the numerical method of computational fluid dynamics. A new Bezier surface modeling method for the entire suction surface and pressure surface of blades is developed, and the multi-island genetic algorithm is directly used for further optimization. Only 32 optimization variables are used to optimize the rotors and stators of the compressor, which greatly overcome the problem of high dimensions, time-consuming calculations, and smooth blade surfaces. After optimization, compared with the original compressor, the peak efficiency is still improved by 0.12%, and the stall margin is increased by 2.69%. The increase in peak efficiency is mainly due to the rotors. Compared with the original compressor, for the second-stage rotor, the adiabatic efficiency is improved by about 0.4%, which is mainly due to the decreases of total pressure losses in the range of above 30% of the span height and 10%–30% of the chord length. Besides, for the original compressor, due to deterioration of the flow field near the tip region of the second-stage stator, the large low-speed region eventually evolves from corner separation into corner stall with three-dimensional space spiral backflow. For the optimized compressor, the main reason for the increased stall margin is that the flow field of the second-stage stator with a span height above 50% is improved, and the separation area and three-dimensional space spiral backflow are reduced.

Under high altitude and low Reynolds number conditions, the aerodynamic performance of compressor cascades deteriorates drastically. In this paper, an optimally designed system combining class-shape-transformation method, S1 surface flow solver and whale optimization algorithm was established to achieve for a controlled diffusion airfoil, called MANGHH. The aim of this work is to improve our understanding of the loss mechanism for the original cascade and optimal cascade under different inflow conditions. The study shows that the total pressure loss of the optimal cascade at an angle of attack of −4°, 0°, and 6° decreases by 55.9%, 16.1%, and 16.3%, respectively, compared with the original controlled diffusion airfoil. The range of the available low loss incidence improves significantly. At different incidences, the optimal cascade moves the blade loading forward compared with that of the original controlled diffusion airfoil while reducing the growth rate of the boundary layer thickness, eliminating a wide range of flow separations. The optimal cascade reduces the total pressure loss mainly by reducing trailing edge mixing loss compared with that of the original controlled diffusion airfoil. Under different inlet Mach number conditions, a laminar separation bubble appears on the suction surface of the original controlled diffusion airfoil. As the inlet Mach number increases, the position of the laminar separation bubble moves slightly upstream, while the length and depth of the laminar separation bubble increase. Fortunately, the total pressure loss of the optimal cascade decreases significantly compared with that of the original controlled diffusion airfoil. Under different incoming turbulence intensity conditions, the total pressure loss of the optimal cascade is always lower than that of the original controlled diffusion airfoil. As the incoming turbulence intensity increases, the total pressure loss of the original controlled diffusion airfoil decreases first and then increases. However, the total pressure loss of the optimal cascade increases with increasing incoming turbulence intensity due to the improvement of the turbulence dissipation capacity.

Aggressive s-shaped compressor transition ducts are important components in the connection between upstream boosters and downstream high-pressure compressors. The flow path is an s-shaped channel with struts and a large radial drop length ratio, which breaks through the limitations of traditional design and has a large aerodynamic loss. Therefore, this paper considers an aggressive s-shaped compressor transition duct in a geared turbofan engine and creatively proposes a method for controlling the flow separation through boundary layer suction. The results show that hub suction reduces the losses of the aggressive s-shape transition duct. As the mass flow rate of hub suction increases, the total pressure loss coefficient decreases and the rate of reduction in the total pressure loss slows down. Combined boundary layer suction reduces the total pressure loss to a greater extent. On the premise that the location of blade suction remains unchanged, the optimal location for the circumferential slot of hub suction is at 20% of the axial chord length of the strut, whereby the total pressure loss coefficient decreases by about 30% compared with the case of no suction. When the mass flow rate of suction is fixed at 3% of the inlet mass flow rate, a distribution of 0.5% from blade suction and 2.5% from hub suction reduces the total pressure loss by 1.6% compared with the case where all 3% comes from hub suction. The distribution of the mass flow rate for combined boundary layer suction has an optimal ratio.

As a degree of freedom in three-dimensional blade design of axial compressors, the sweep technique significantly affects aerodynamic performance of axial compressors. In this paper, the influence of backward sweep rotor configurations on aerodynamic performance of a 1.5-stage highly loaded axial compressor is studied by numerical simulation. The aim of this work is to improve