Investigation On Integrated Cooling Performance And Flow Resistance Characteristics Of Vane

Development of high thrust weight ratio engine requires to improve the turbine inlet temperature and reduce air consumption at the same time.Carry out research on new efficient cooling technology and tap cooling potential of cooling method,also design the blade cooling structure accurately is essential.It is necessary to evaluate the comprehensive cooling effect of turbine guide vane by the method of verification experiment.Based on the purpose of optimizing the cooling structure and cooling capacity distribution of turbine guide vane,a cascade wind tunnel test rig is built.The infrared temperature measure technology is applied to study the comprehensive cooling characteristics of original structure and other 4 modified structure blades under the mass flow ratios of 0.059~0.118.Analyzing the comprehensive cooling efficiency and the coolant flow resistance characteristics of the original structure blade in detail,the two-dimensional distributions of the comprehensive cooling efficiency on the vane and the spanwise averaged and the surface averaged comprehensive cooling efficiency were obtained.The variation rule of conversion flow coefficient with the pressure ratio was obtained.In the same conditions,the original and modified structure blade of high-precision parameter comparison test was carried out.The highest comprehensive cooling efficiency of the composite cooling structure was obtained.In order to reduce the influence of errors in the two test of leaf replacement,repeated stability test is carried out to ensure that the errors of comprehensive cooling effect of each repeated test leaves is less than 2%.The results show that the composite cooling structure with the internal impact joint and gas film is the best,and the original structure comes second.When the leading edge is cooled without gas film,the cooling effect of the leading edge area is very low,and the flow resistance of the cooling air inside the cavity increases.The influence of the structural change of the inner impact hole on the comprehensive cooling effect and the cold air flow resistance can be ignored.Comparing with the dustpan type gas film hole in the suction side,the structure of hole slot film hole can reduce the comprehensive cooling effect’s descending gradient which caused by the difference of the cooling mode on the suction side.The effect of flow ratio change on the average comprehensive cooling effect in the leading edge and pressure side is obviously greater than that of the suction side.The maximum average comprehensive cooling effect of the spanwise directions of the leading edge is near the stagnation line.When the cooling air flow rate gradually decreases,the area where the comprehensive cooling effect decreases obviously first appears in the 0.04~0.1 arc length range on the side of the leading edge near the pressure side,and the comprehensive cooling effect of the blade direction is manifested as decreasing along the inflow direction of the inner cavity cooling air.The comprehensive cooling effect of the blade is the highest when ω=0.078,and the comprehensive cooling effect of the blade is reduces obviously when ω< 0.049.