Thermal Performance of a Cooled Vane in a Transonic, Annular Cascade

Abstract

Gas turbines continue to be an integral technology in both the energy and transportation sectors due to their relatively low cost of energy (COE) and high thrust-to-weight ratio. Extended mission profile requirements necessitate more power without increasing the engine size, typically achieved by increasing the turbine inlet temperatures. Further, efficient engines are needed to minimize COE and reduce emissions. In the first stage of the turbine, the gas temperature is the hottest and the flow is often transonic. Enabling increased power output, higher efficiency, and reliable operation depends on optimized cooling techniques in this region. Film cooling is a common insulating technique for protecting gas turbine blades and vanes, complementing thermal barrier coatings and internal cooling. To quantify the reduction in heat transfer, binary pressure sensitive paint (BPSP) was used in three discrete studies to measure the film cooling effectiveness and the heat transfer coefficient. A novel experimental technique was developed to allow measurement of the heat transfer coefficient with BPSP on a leading-edge cylinder at low speeds, then applied on the endwall of a transonic annular cascade with upstream and passage film cooling holes. Based on measurements of the film cooling effectiveness at an exit Mach number of 0.7 and 0.9 and a density ratio of 2, a minimum coolant-to-mainstream mass flow ratio (MFR) of 0.75% or 1% through the passage holes is recommended to prevent mainstream gas ingestion and jet liftoff. Based on the measurements of both film cooling effectiveness and heat transfer coefficient measurements at an exit Mach number of 0.9 and a density ratio of 1, an optimum upstream MFR of 0.75% is recommended. This study has provided useful information for gas turbine designers to optimize film cooling. It has also demonstrated a novel experimental technique to fully quantify film cooling heat transfer.