Experimental and Numerical Investigations of a Transonic, High Turning Turbine Cascade With a Divergent Endwall

Abstract

The paper presents detailed measurements of midspan total pressure loss, secondary flow field, static pressure measurements on airfoil surface at midspan, near hub and near the end walls in a transonic turbine airfoil cascade. Numerous low-speed experimental studies have been carried out to investigate the performance of turbine cascades. Profile and secondary loss correlations have been developed and improved over the years to include the induced incidence and leading edge geometry and to reflect recent trends in turbine design. All of the above investigations have resulted in better understanding of flow field in turbine passages. However, there is still insufficient data on the performance of turbine blades with high turning angles operating at varying incidences angles at transonic Mach numbers. In the present study, measurements were made at +10, 0 and -10 degree incidence angles for a high turning turbine airfoil with 127 degree turning. The exit Mach numbers were varied within a range from 0.6 to 1.1. Additionally, the exit span is increased relative to the inlet span resulting in one end wall diverging from inlet to exit at 13 degree angle. This was done in order to obtain a ratio of inlet Mach number to exit Mach number which is representative to that encountered in real engine and simulates the blade and near end wall loading that is seen in an engine. 3D viscous compressible CFD analysis was carried out in order to compare the results with experimentally obtained values and to further investigate the design and off-design flow characteristics of the airfoil under study. All aerodynamic measurements were compared with CFD analysis and a reasonably good match was observed. INTRODUCTIONWith present day stringent norms set by the government for clean and efficient engines and the ever-growing competition in the market, gas turbine manufactures need to optimize every component of the engine. The need for low cost, compact and fuel efficient engines are apparent. Improvements can be made by designing blades such that minimum losses are generated while making sure that the loading on the blade is maximized in order to reduce the number of airfoils in each stage. The designer also needs to ensure that the turbine efficiency is maintained at off-design conditions. Almost a third of the total losses in turbines are due to end wall losses. The thickness of the upstream boundary layer as well as the blade turning angle influences the strength of the secondary flow observed near the end walls. The secondary flow results into stagnation pressure loss which accounts for a considerable portion of the total stagnation pressure loss occurring in a turbine passage.