Aero-thermal redesign of a high pressure turbine nozzle guide vane

Abstract

The current article presents conceptual, preliminary and detailed aero-thermal redesign of a typical high pressure turbine nozzle guide vane. Design targets are lower coolant consumption, reduced manufacturing costs and improved durability. These goals are sought by 25% reduction in vane count number and lower number of airfoils per segment. Design challenges such as higher airfoil loading, associate aerodynamic losses and higher thermal loads are discussed. In order to maximize coolant flow reduction and avoid higher aerodynamic losses, airfoil Mach distribution is carefully controlled. There has been an effort to limit design changes so that the proven design features of the original vane are used as much as possible. Accordingly, the same cooling concept is used with minor modifications of the internal structures in order to achieve desired coolant flow and internal heat transfer distribution. Platforms of the new design are quite similar to the original one except for cooling holes and application of thermal barrier coating (TBC). Detailed aerodynamics/heat transfer simulations reveals that the reduced trailing edge (T.E.) blockage and skin friction dominated the negative effect of increased secondary losses. As a result the reduced design performs acceptable in terms of total pressure loss and improving stage efficiency for a wide range of varying pressure ratio. Moreover, more than 20% cooling mass flow can be saved; while maximum and average metal temperatures as well as cross sectional temperature gradients have not been changed much.