Investigation On Streamline Curvature Method For Multistage Compressor With Considering Viscosity And Spanwise Mixing Effects

High load multi-stage compressor is the most complex part of aero engine, the technological difficulty is how to get the distribution of aerodynamic parameters along the leading and trailing edge of blades under all working conditions and the whole engine performance quickly, completely and accurately. Although nowadays the3D CFD technology is becoming more and more mature, through flow calculation is still the core technology in multi-stage compressor design based on its quick calculating speed and explicit orientation of result, and was updated continuously. This thesis based on this background, focus on increasing the applicable scope and accuracy of the code, considering the effect of boundary layer and spanwise mixing, researching about the high efficiency through flow performance calculation method. My work includes the following aspects:1. Research a lot of papers about the compressor through flow performance calculation. Summarize the latest research progress and research direction of compressor through flow calculation in inland and abroad.2. An uniform streamline curvature equation for both axial and centrifugal compressor is discussed in this article, at the same time, consider the tip region viscosity for the effect of boundary layer and consider spanwise mixing.3. Make a conclusion of the attack angle, deviation angle, loss and blocking empirical model by considering their available condition.4. Develop a new through flow performance prediction code for axial compressor from the very beginning. Based on the requirement of modularity, use OOP(object-oriented programming) and modern coding standards. By this code, the performances of NASA Rotor37, Rotor67, etc. were calculated, and the results show that the new code is effective.