Aerodynamic Design Of High Performance Two Stage High-Pressure-Ratio Compressor

Compressor is one of the main components of aeroengine,and its design level has a critical influence on the overall performance of the aeroengine.The compressor requires high performance not only at design point,but also at off-design point.The design of multistage compressor faces many challenges,such as high stage pressure ratio,narrow stable working range and difficulty in matching between stages.The matching performance can improve by increasing the surge margin of each stage.The performance at design point and surge margin are taken into account in the paper.The optimization method is proposed to design the blade with the performance at design point as target,and then modify the rotor/stator blade geometry manually to explore the method to improve compressor surge margin,so as to design a two stage high-pressure-ratio compressor with high performance and wide surge margin.In order to improve the accuracy of flow field calculation,grid distribution was discussed in detail.NUMECA software was used to study the influence of yplus,grid point distribution in the azimuthal and spanwise direction as well as grid size on CFD results using transonic fan NASA Rotor67 and supersonic compressor NASA Rotor37.The numerical result was compared with experimental data to determine the appropriate mesh size and grid point distribution.The S-A turbulence model,y+?1,and approximate 500,000 grid points can ensure the reliability of the design point pressure ratio,isentropic efficiency and surge margin.The optimization method was used to design the blade with the performance at design point as the objective.The initial two-dimensional rotary profile was designed according to the S2 throughflow design results(the rotor profile was designed using supersonic profile parametric method base on unique incidence theory),and then improved using two-dimensional profile optimization program by modifying profile line,installation angle and chord length automatically.The optimized blade profile was stacked radially,and then three-dimensional optimization design was carried out for variables such as camber line,thickness,chord length,installation angle,stacking line sweep and lean,meridional channel.The optimization of two-dimensional profile and three-dimensional blade both adopted automatic optimization design program based on genetic algorithm and parametric method.The first-stage rotor was designed and optimized referring to NASA Rotor37.Total pressure ratio of optimized rotor at design point is 2.053,slightly less than design object 2.106,but its isentropic efficiency can reach 0.9247,which is much better than NASA Rotor37.The design method of stator blade is similar as that of the rotor blade,but three-dimensional flow field calculation of the stator was carried out in the stage environment to consider the influence of the upstream rotor.The total pressure ratio of first-stage at design point is 1.975,the isentropic efficiency is 0.8942,and the surge margin is 18.33% after the optimization of the stator.Methods to improve compressor surge margin was explored on the basis of optimized blade.The rotor stalls when shock wave pushed out of the passage at the tip.The surge margin of the single rotor improved from 7.10% to 18.29% if measures such as forward sweep,increasing blade tip solidity are taken to control the shock position in the passage.The predicted surge margin of first-stage improved from 18% to above 30% if methods such as forward sweep,tangential lean,modification of blade number and geometric inlet angle are adopted for the stator.Forward swept stator can reduce static pressure at the exit of rotor and decrease large separation at near stall point,and increasing the geometric inlet angle of the stator can reduced the attack angle at near stall point,which can improve surge margin obviously.The similar method was used for the design of the second-stage,in which three-dimensional flow field calculation of second-stage rotor/stator were all carried out in the stage environment.The two stage compressor calculation together was carried out and modification of the second-stage stator geometry inlet angle to expand the surge margin.The total pressure ratio of the two-stage high-pressure compressor at design point is 3.517,the isentropic efficiency is 0.8736,and the surge margin is 17.53%.The two compressor has well performance at design point and wide surge margin at the same time.