Multi-objective Optimization Of Film Cooling Of Turbine Blades Based On Source Terms Method

As one of the most advanced energy conversion machines in the world,gas turbine plays an important strategic role in the field of power machinery.With the continuous progress of society,gas turbines are also developing towards high performance,high parameters and light weight.The high performance and high parameters mean that the temperature in front of the gas turbine is increasing continuously,which brings extremely high heat load to the turbine blades.As a simple and efficient cooling method,film cooling is widely used in the cooling of turbine blades.Combined with experimental design(DOE),source terms numerical simulation(SCFD)and response surface optimization method(RS),this paper aims at multi-objective optimization of aerodynamic heat transfer performance of a turbine blade tip and blade surface,aiming at obtaining a film cooling scheme with excellent gas and heat performance.The research content includes the following three aspects:(1)Verification of source terms numerical simulation accuracy.In order to reduce the calculation cost in the optimization process,this paper adopts a CFD technology based on source terms.The method does not need to build a real film hole structure and generate a body-fitted grid,but only needs to build a jet source terms domain near the film hole,then define the material porosity at the grid point,and introduce the jet source terms function related to the material porosity into the control equation.The aerodynamic flow field of a turbine cascade with an airfoil crown was calculated by using the numerical simulation method based on the source terms.Then,based on the source terms of the airfoil crown,the jet source terms of the sealing teeth and the tip were added to study the accuracy of the source terms method in the performance calculation of the jet cascade with sealing teeth and tip.Compared with the numerical simulation results based on body-fitted grid(i.e.,real structure),it is found that the source terms method can accurately evaluate the influence of airfoil crown,sealing teeth and tip injection on the aerodynamic performance of turbine cascade.(2)Multi-objective optimization of film cooling on turbine blade tip.Adopting DOE experimental design and response surface optimization method,firstly,the optimized design variable,i.e.,axial geometric position information of tip injection holes,is sampled,and the objective function,i.e.,aerodynamic performance and heat transfer performance of cascade,is evaluated by using source terms numerical simulation.after DOE experimental calculation is completed,response surface fitting and optimization are carried out on the obtained sample point information,in order to obtain a tip injection hole arrangement scheme with excellent gas and heat performance.(3)Multi-objective optimization of film cooling on turbine blade surface.In order to reasonably arrange the position of the gas injection holes on the turbine blade surface,the same optimization method was used to optimize the aerodynamic heat transfer performance of the axial geometric position of the gas film holes on the blade pressure surface.The geometric reconstruction of the optimized scheme was carried out,and the numerical verification was carried out by body-fitted grid.The results showed that the optimized gas film holes increased the average heat flux on the blade surface by 15.62% under the same cold air flow rate,and the cooling effect was obviously improved.