Aerodynamic Performance of High Turning Airfoils and the Effect of Endwall Contouring on Turbine Performance

Gas turbine companies are always focused on reducing capital costs and increasing overall efficiency. There are numerous advantages in reducing the number of airfoils per stage in the turbine section. While increased airfoil loading offers great advantages like low cost and weight, they also result in increased aerodynamic losses and associated issues. The strength of secondary flows is influenced by the upstream boundary layer thickness as well as the overall flow turning angle through the blade row. Secondary flows result in stagnation pressure loss which accounts for a considerable portion of the total stagnation pressure loss occurring in a turbine passage. A turbine designer strives to minimize these aerodynamic losses through design changes and geometrical effects. Performance of airfoils with varying loading levels and turning angles at transonic flow conditions are investigated in this study. The pressure difference between the pressure side and suction side of an airfoil gives an indication of the loading level of that airfoil. Secondary loss generation and the 3D flow near the endwalls of turbine blades are studied in detail. Detailed aerodynamic loss measurements, both in the pitchwise as well as spanwise directions, are conducted at 0.1 axial chord and 1.0 axial chord locations downstream of the trailing edge. Static pressure measurements on the airfoil surface and endwall pressure measurements were carried out in addition to downstream loss measurements. The application of endwall contouring to reduce secondary losses is investigated to try and understand when contouring can be beneficial. A detailed study was conducted on the effectiveness of endwall contouring on two different blades with varying airfoil spacing. Heat transfer experiments on the endwall were also conducted to determine the effect of endwall contouring on surface heat transfer distributions. Heat transfer behavior has significant effect on the cooling flow needs and associated aerodynamic problems of coolant-mainstream mixing.

One of the primary objectives of this study is to provide data under transonic conditions that can be used to confirm/refine loss predictions for the effect of various Mach numbers and gas turning. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. A published experimental study on the effect of end wall contouring on such high turning blades at high exit Mach numbers is not available in open literature. Hence, the need to understand the parametric effects of endwall contouring on aerodynamic and heat transfer performance under these conditions.