Fast Three-Dimensional Numerical Study Of Multi-Stage Compressor

With the increase of pressure ratio and aerodynamic load for modern multi-stages axial-flow compressors, the internal flowfield in compressors becomes more and more complicated. To obtain a better insight of the real flows in compressors and hence improve the aerodynamic performances of compressors, numerical simulation method has to be adopted because of its capability of investigating on the complicated flowfield.To meet the requirements for 3D viscous flow computational softwares in compressor design, a fast full 3D viscous numerical method for the computation of overall aerodynamic performance in multi-stages axial-flow compressor was introduced. Effective wall functions were employed to reduce the overall computational effort. Combined with compressible Reynolds-averaged Navier-Stokes equations, a numerical technique based on a new Lower-Upper implicit scheme and a fourth-order high MUSCLTVD scheme has been developed.The numerical simulations have been performed for NASA Rotor 67 transonic fan rotor and Rotor37 transonic fan rotor at several operating condition. The good agreements between the computational and experimental results verified the validation and precision of the fast numerical method. The computational results were evaluated by comparing with the results obtained from the normal precise full 3D numerical method and the results extracted from domestic and overseas literatures. It can been known by comparisons that the fast numerical method presented in this paper can get overall performance of 3D viscous flowfield rapidly and, the precision of the computational results should be acceptable in engineering application. In view of the adoption of the simple wall functions, however, the present numerical method might result in large numerical errors for the simulation of compressors in which strong shock/boundary interactions exist.The personal computer network parallel computational technique and mixing-plane model were adopted to extend the present single-blade-row algorithm to multi-blade-row simulations. The advantages, defects and realizing program of the personal computer network parallel computational technique were thoroughly studied and the computational efficiency of the parallel technique was displayed. The advantages, defects, selection of transferring data and realizing program of the mixing-plane model were particularly discussed.Finally, an interface to control the calculation was designed. And the compressor of a turbofan engine and the first five stages of a nine-stage compressor werecalculated to prove the reliability of this fast numerical simulation code. As a result, it was confirmed that the method was accurate and reliable for some subsonic and slight transonic compressors, but should be careful for the high transonic and supersonic compressors. So the suitable application range for the code designed in this paper should be clear with the help of the above-mentioned discussions. At last, some suggestions to improve this code were presented in order to extend the usable range of this code.