Investigation Of Transonic Compressor Rotor For Three-dimensional Aerodynamic Design And Experiment

Gas turbine,as an advanced and complex high technology product,represents the comprehensive developing level of multi disciplines and engineering technology fields.The gas turbine industry,which new material and new technologies are rolled into one,has a very prominent strategic position in a country and is an important symbol of the national high technology level and technical strength.As one of the three core components of a gas turbine,the compressor's performance has tremendous influence on the gas turbine's performance which can be improved by the optimization design of the compressor.The number of stages of a compressor,the axial size and weight of a gas turbine can be reduced by improving the compressor single-stage load and increasing the rotor' s circumferential speed.So,to explore high load transonic compressor has become the development direction of the compressor.However,there exists contradiction between higher stage pressure ratio of the compressor and high efficiency and high stability margin.Therefore,based on the understanding of the internal flow and an elaborate optimization design,analyzing the loss sources and the key factors which affect the flow stability of the transonic compressors,only can we obtain the superior performance of a high load transonic compressor.The single stage transonic compressor rotor is the research object in this paper.The changes of the internal flow field parameters of the rotor under different working conditions,and the loss sources and the key factors which affect the flow stability of the transonic compressor are analyzed.The results show that there exists low efficiency area near the root,middle and the top of the rotor blade.At the near stall condition,the tip leakage vortex and the shock wave/boundary layer interactions lead to the break of the tip leakage vortex,and the low energy fluid centralizes here and blocks the flow channel.It is the main reason to cause the compressor to unstable condition.In order to improve the transonic compressor pressure ratio,efficiency and stable working range,the optimization design is carried out by combining the methods of the genetic algorithm and artificial neural network at same flow rate.The root,5%blade height,50%blade height,95%blade height of the rotor blade are selected to parameterization profiling with no change to the hub,case line and blade stacking,only change to the blade stagger,the wedge angle of the trailing edge and the suction surface control parameters.The optimization blade profile OPT1 and OPT2 with better performance are obtained finally by means of the Latin hypercube sample selection method to get the sample database,and the artificial neural network to autonomic learning and training.The results show that compared with the ORI rotor,the OPT1's efficiency and pressure ratio can be increased 0.91%and 0.52%,and the OPT2's efficiency and pressure ratio 0.80%and 1.04%.Through the internal flow field analysis of the prototype and optimization blade profile,the position of the shock waves at the blade top can be changed and move to the trailing edge by reducing the stagger at 50%blade height and increasing the stagger at 95%blade height and the blade torsional degree.Combined with the adjustment the wedge angle of the leading and trailing edge,the blade load distribution of the compressor rotor can effectively controlled and the rotor pressure ratio and efficiency can be improved.Then,the plane cascade experiments with the blade profiles of the prototype and the OPT1's at 5%blade height,50%blade height and 95%blade height section.The influence rules of blade profile change on the flow parameters and the flow parameters of the incidence characteristics at different blade height section of the prototype and the OPT1 are explored and researched.The results show that,at 5%blade height,the incidence of minimum loss moves to negative incidence direction,and the control of the OPT1 to the corner separation is superior to the prototype.At 50%blade height,the incidence loss characteristics of the two are similar,but the OPT1's stable range of the incidence is higher than that of the prototype.The minimum loss incidence of the OPT1 moves to positive incidence direction compared to the prototype,and the stable range of the incidence increases,but the control to the corner separation is weaker than the prototype at 95%blade height.In the range of the Mach numbers,the total pressure loss coefficient of the OPT1 is less than that of the prototype.Finally,in order to use the experimental results of the plane cascades to the three dimensional blade design of the rotor,the experimental and numerical calculation results of the plane cascades are compared with the numerical calculation results of the three dimensional rotor,and the influence of the aerodynamic performance of three dimensional rotor profiling is analyzed under different rotational speed.Under 0.6 design rotational speed,the total pressure loss coefficients,the outlet flow angles and blade surface pressure coefficients at 5%and 50%blade height are coincident between the plane cascade and the three dimensional rotor basically.Influenced by the tip clearance leakage flow,the total pressure loss of the rotor increases and the blade suction surface pressure raises due to the interaction of the leakage flow and the blade surface boundary layer at 95%blade height.In contrast to the outlet Mach numbers,the Mach numbers of the three dimensional rotor are higher than that of plane cascades which shows that the three dimensional profiling can reduce the kinetic energy loss.Under 1.0 design rotational speed,the three dimensional profiling of the rotor has an important influence on the cascade performance,and the influence on the distribution rules of the total pressure loss,outlet Mach numbers,and outlet flow angle are similar to that of 0.6 design rotational speed.The difference of flow field parameters between the three dimensional rotor and the plane cascade at 5%blade height is mainly caused by the flow channel contraction and centrifugal force which can weaken the reverse pressure gradient and the intensity of shock waves,change the shock waves to compression waves,result in loss reduction of the three dimensional rotor root.The main difference between the three dimensional rotor and the plane cascade in which the flow channel contraction is coincident basically at 50%blade height is the centrifugal force.The radial separation vortex is formed and moves to the tip direction due to the interaction of the centrifugal force,shock wave and boundary layer above 30%blade height in the rotor.The exist of the radial separation vortex makes the flow velocity reduce,the static pressure rise and causes the static pressure on the blade surface higher than that of the plane cascade.The flow is the most complex and the flow loss is serious in the rotor at 95%blade height.Because the common interaction of the three dimensional tip clearance leakage flow,the flow channel contraction and centrifugal force,the total pressure loss coefficient of the three-dimensional rotor is much higher than that of the plane cascade,but the blade surface pressure distribution is not affected more before 50%blade chord.