Experiment And Numerical Investigation Of Transonic Compressor Cascade

Compressor is an important part of aircraft engine, and it will influence the performance of the overall engine performance directly. Therefore, it is necessary to make a further study on the complex flow in the compressor, and verify the design of the compressor. In this paper, the method of wind tunnel test and numerical simulation was combined to make a detailed analysis on the cascade flow field. First of all, a set of compressor cascade was designed and manufactured, the transonic wind tunnel test was conducted at several incidence and inlet mach number, and the flow field data was obtained for different test conditions. Under different flow conditions, the change rule of parameters such as total pressure loss coefficient, static pressure ratio, isentropic mach number, surface pressure coefficient and deviation angle was obtained after processing the experiment results.Secondly, under the same conditions as the experiment, the experiment cascade was simulated and then compared with the experiment results by using ANSYS CFX software. The Comparison verifies the accuracy and reliability of the calculation model. Meanwhile, the blade profile will be used as an original profile, and the result of numerical simulation will be used as reference for comparison calculation.Thirdly, the blade profile parameterized calculation method was integrated in this paper. A set of compressor blade modeling two-dimensional design platform was developed with Visual Studio. NET. The software can be used to modify the blade profile directionally based on the blade profile parameterized calculation method. Two groups of new blade profile were obtained respectively by modifying the front oval short-long axis ratio and the maximum thickness location parameter based on the original blade profile.Finally, the new type of cascade was simulated by ANSYS CFX software under the same calculation conditions and compared with the results of numerical simulation of original blade profile. The comparison shows that the total pressure loss will increase and the pressure distribution of blade leading edge will be affected as the leading edge oval short-long axis ratio increases; The minimum pressure point of blade suction surface will move forward and inverse pressure gradient length will increase as the position of maximum thickness moves forward.