Endwall component for a turbine stage of a gas turbine engine

#The present invention relates to a component of a turbine stage of a gas turbine engine, the component forming an endwall for the working gas annulus of the stage.

#With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate-pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.

#The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.

#Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:

#FIG. 1 shows a schematic longitudinal cross-section through a ducted fan gas turbine engine;

#FIG. 2 shows an isometric view of a typical single stage cooled turbine;

#FIG. 4 shows a schematic cross-sectional view through a high-pressure turbine shroud segment according to a first embodiment. The shroud segment has an endwall which forms a gas-washed surface for the working gas annulus of an engine. Internal plena 41 are formed behind the endwall, the plena containing a flow of cooling air introduced into the plena through feed holes 42. In FIG. 4 two plena are shown, but the number could be as low as one or perhaps as high as five or six. A plurality of exhaust holes 43 traverse the endwall, each hole has an entrance 44 which receives cooling air from the plena and an exit 46 at the gas-washed surface from which the cooling air effuses to form a cooling layer over the gas-washed surface.

#Each exhaust hole 43 expands in flow cross-sectional area from its entrance 44 to a maximum area at an intermediate position 45, and then contracts in flow cross-sectional area to its exit 46. The flow cross-sectional area at the intermediate position can be greater than the flow cross-sectional area at the entrance and/or the exit by a factor of at least 1.5, and preferably by a factor of at least 2 or 4.

#There is a pulse of high pressure in the hot working gas as each turbine blade passes over the shroud segment. Due to their increased flow cross-sectional area at the intermediate position 45, the exhaust holes 43 have high internal volumes relative to conventional straight exhaust holes. Accordingly, flow of ingested hot gas through each exhaust hole 43 has to expand at the intermediate position. This in turn produces an increased pressure loss when the hot gas enters the exhaust hole. This pressure loss helps to retain the ingested hot gas in the exhaust holes for a given pressure of the cooling air in the plena. That is, the cooling air in the plena is maintained at a pressure which prevents hot gas ingestion into the plena at the peak of each pressure pulse, but by adopting exhaust holes of the type shown in FIG. 4 that pressure can be reduced, leading to consequent improvements in engine efficiency. Some hot gas ingestion into the exhaust holes occurs, but as long as the hot gas is prevented from mixing with the cooling gas in the plena, that hot gas is simply ejected from the holes after the peak of the pressure pulse is passed.