Optimal Design Of Axial Compressor Blade Aerodynamic Performance

Gas turbine plays an extremely important role in military and civil fields.As one of the three core components of gas turbine,compressor is always in the key position of gas turbine performance research.As the main part of the compressor,the blade’s external geometry and aerodynamic characteristics play a decisive role in many factors affecting the adiabatic efficiency of the compressor.Therefore,it is of practical significance to carry out blade optimization design.In this paper,the 2.5-stage transonic axial compressor in the laboratory is taken as the research object to carry out the aerodynamic performance optimization design of compressor blades.The main research contents are as follows:1)The flow field of the compressor under different working conditions was numerically simulated,the compressor characteristic curve was drawn,and the accuracy of the numerical simulation was verified by comparing the experimental results,so as to provide a reliable numerical simulation scheme for the aerodynamic performance optimization of the compressor blade.The characteristics of flow loss in the flow passage of compressor are analyzed to provide reference for the optimization of aerodynamic performance of compressor blades.2)Based on the analysis of the characteristics of different fitting curves,different flow surfaces and radial distribution modes,and different stacking modes of blades,the Bessel curve and b-spline curve were selected for parametric fitting of compressor blades,so as to provide high-precision parametric profile of compressor blades for aerodynamic performance optimization of compressor blades.3)Combining the artificial neural network and genetic algorithm,12 parameters that can control the blade bending and sweep characteristics were selected from all the control parameters of the fitted blade to be adjusted.Under the design condition,the optimal design and calculation were carried out to improve the adiabatic efficiency,and the final optimized blade was obtained through 30 steps of iteration.The results show that the influence of the second stage blade on the flow field of the compressor is higher than that of the first stage blade.Under the condition that the pressure ratio is basically unchanged,the adiabatic efficiency of the compressor after the two-stage dynamic blade and its matching stator blade profile optimization is improved by 0.82%,and the trailing edge vortex of the two-stage stationary blade tip disappears.