Novel Turbine Rotor Shroud Film-Cooling Design and Validation: Part 1

Abstract

This paper is part one of a two part paper which considers a shroud film-cooling system designed using a two-dimensional approach. Heat transfer to rotor-casings has reached levels that are causing in-service difficulties to be experienced. Future designs are likely to need to employ film-cooling of some form. There is currently very little information available for film-cooling on shroudless turbine rotor-casing liners. Heat transfer literature on uncooled configurations is not extensive and in particular, spatially-detailed, time-accurate data are rare. This paper describes the aero-thermodynamic design and validation of a rotor casing film-cooling system for a transonic, high-pressure shroudless turbine stage. The design was carried out using a boundary layer code with the film-cooling hole geometry representative of an engine configuration and, has been subjected to mechanical constraints similar to those for an engine component. The design consists of two double rows of cooling holes and two ‘cooling-hole’ shape configurations, cylindrical and fan shaped. The design was tested in the QinetiQ short duration turbine test facility (TTF). Measurements taken include casing heat transfer using thin film gauges and stage exit total pressure, Mach number and flow angle using a three-hole pressure probe. Results showed that while the cooling produced a reduction in the heat transfer rate close to the injection point, the film was stripped off the casing and entrained in nozzle guide vane secondary and rotor overtip flow, where it was transported spanwise towards the hub in the rotor passage. Using the results obtained from this deign a second cooling design was carried out, using a three-dimensional approach this gave significantly better cooling performance. The thee-dimensional design and validation is reported in GT2009-60246 as part 2 of this paper.