Research Of Full Three Dimensional Viscous Inverse Design Method For Multi-stage Axial Compressor

Nowadays,the existing axial compressor aerodynamic design system is based on the Qusi-3D throughflow calculation.The Computational Fluid Dynamics is regarded as the flow solver during the design process.In recent years,engineers were trying to develop a new axial compressor design method which is based on the Full-3D CFD.In order to make CFD used in the axial compressor aerodynamic design process more intuitively,the inverse design method is continuously developed.The axial compressor inverse design theory and method which is based on solving the Navier-Stokes equations are researched in this paper.The computer code which contains both direct mode and inverse mode is developed.The application of inverse method is studied based on different types of axial compressors.The Navier-Stokes equations of Cartesian are used to calculate the flow field of axial compressor.During the calculation,the cell centered Finite Volume method is implemented.Both the JST central scheme and Steger-Warming upwind scheme are applied to calculate the convection flux on the control volume surface.The explicit hybrid Runge-Kutta scheme and implicit LU-SGS scheme are used for the time stepping.The turbulence viscous is simulated by Baldwin-Lomax turbulence model.Two acceleration techniques are applied to improve the solving efficiency.In order to validate the accuracy of current axial compressor flow field solver,several different types of axial compressor are used as the test case.The calculation results and experiment results are compared and analyzed.The calculation result shows good agreement with the experiment result,which demonstrate the efficiency and accuracy of axial compressor flow solver.Two different axial compressor inverse design method are developed based on the inverse design theory,the loading-camber inverse design method and pressure-profile inverse design method.The two inverse design method using loading or pressure distribution on the blade surface as the design input condition,and the blade camber or profile is chosen as the design object.The core problem of the inverse design method is the inverse boundary condition.Different inverse boundary conditions are developed corresponding to the two different inverse design methods.The blade geometry updating method and the inverse design process are introduced in detail.The dynamic meshing technical is used to update the computational mesh after the blade geometry is changed.In order to validate the effectiveness of present inverse method,the blade recovery test is carried out.The result shows that the inversedesigned blade matched the target blade both in aerodynamic parameters distribution and blade geometry,which demonstrate the accuracy of present inverse methods.The application of axial compressor inverse design method is researched based on the engineering practice.In order to eliminate the flow separation in the subsonic flow,the flow separation region near the hub of Rotor 67 is chosen as the research object.The flow separation is controlled by adjusting the loading distribution on the blade surface.Multi-Rows inverse design method is studied.The loading distribution on the both rotor and stator blade of DLR single transonic compressor stage is modified based on the flow feature of each blade.Using the modified loading distribution as design input,the inverse designed compressor aerodynamic performance is improved.The inverse design method is applied to aerodynamic matching for multi-stages compressor.The 4 stage compressor is chosen as the test case and the relationship between blade exit flow swirl and loading distribution is established.The loading distribution is adjusted automatically during the inverse design process until the specified blade exit flow swirl is satisfied,which improve the aerodynamic matching between axial compressor stages.In order to control the shock wave strength in the blade passage,the static pressure distribution on the blade surface near the tip of Rotor 67.The pressure gradient is decreased near the shock wave and the shock wave strength and loss is reduced.The similar method is used to control the flow separation which is caused by the shock wave/boundary layer interaction,the flow separation is eliminated by reduce the positive pressure gradient near the blade surface.In order to research the application of inverse design method more particular,the stator of Stage 35 is redesigned by using the concept of Controlled Diffusion Airfoil.The isentropic Mach is utilized to build the relationship between the flow speed and static pressure on the blade surface.The flow structure in the blade passage is successfully controlled by modifying the pressure distribution on the blade surface.