Numerical Study Of Transonic Compressor Plane Cascade Performance Affected By Leading Edge Design

Gas turbine is of great strategic significance to the national defense industry and the national economy.With the development of compressor towards high-lift and super/transonic speed,the design technology of high performance compressor blade has become one of the hot research fields.As the starting position of the blade surface boundary layer,the change of the leading-edge geometry will affect the flow field distribution.The reasonable leading edge treatment can effectively control the flow separation,improve the flow field structure and the working performance of the compressor.In this paper,the 5%typical section of the rotor DMU37 is taken as the research object.Calibration of numerical simulation software FLUENT by experimental data is used to conduct a detailed numerical study of the prototype and the design of the cascade.By analyzing the total pressure loss coefficient,the distribution of static pressure coefficient and Mach number,explore the influence of the blade leading edge design on the flow characteristics and the loss mechanism of the transonic compressor,seek a useful design direction,lay the foundation for further front flow control technology and necessary experiment.Firstly,the grid independent verification is carried out,and based on the comparison with the experimental data,the suitable turbulence model is chosen to ensure the accuracy and reliability of the calculation method.Then,the numerical simulation of the cascade flow field is carried out,the results show that with the increase of the Mach number,the cascade expansion ability is enhanced and the flow loss increases.The flow condition is better under the negative incidences,with the increase of the incidence angle,the position of the leading edge saddle is advanced and offset,the inverse pressure gradient is enhanced,the separation region of the end angle area and the loss increases.Finally,in the premise of not changing the prototype blade chord,a blade radial parameter modeling method is proposed,compared and analyzed the flow field of the cascade under different flow conditions,the results show that the total pressure loss of leading edge designed increases,outlet pressure ratio the cascade power capability improve.The end wall region of low-energy fluid group decreases,leading adverse pressure gradient and the corner separation becomes weak,the turning ability of airflow increased,flow loss becomes smaller,Among them,the incidence angle of designed cascade can be calculated from +4 degrees to +6 degrees,but the cutting depth should not be too large.On this basis,the change of the radial cutting height and the retention length of the end wall does not affect the total pressure loss optimization of turning point.Larger cutting height,smaller length of end walls to retain,the optimization effect of the flow field is preferable.Compared with the higher Mach number,the different cutting programs have better effect on the lower Mach number.