Abstract: A method of manufacturing a coated turbine vane (34) comprises manufacturing a turbine vane (34) having a platform (44) and an aerofoil (42) extending from the platform (44), a curved transition (60) connects the platform (44) to the aerofoil (42) and a recess (64) is provided in the curved transition (60) from the platform (44) to the aerofoil (42). A bond coating (70) is deposited on the platform (44), the aerofoil (42), the curved transition (60) and the recess (64). A ceramic thermal barrier coating (72) is deposited on the platform (44), the recess (64) and the curved transition (60) by plasma spraying. The recess (64) reduces the size of the step due to the ceramic thermal barrier coating (72) and hence improves the aerodynamics of the turbine vane (34).

Abstract: A method of manufacturing a coated turbine vane (34) comprises manufacturing a turbine vane (34) having a platform (44) and an aerofoil (42) extending from the platform (44), a curved transition (60) connects the platform (44) to the aerofoil (42) and a recess (64) is provided in the curved transition (60) from the platform (44) to the aerofoil (42). A bond coating (70) is deposited on the platform (44), the aerofoil (42), the curved transition (60) and the recess (64). A ceramic thermal barrier coating (72) is deposited on the platform (44), the recess (64) and the curved transition (60) by plasma spraying. The recess (64) reduces the size of the step due to the ceramic thermal barrier coating (72) and hence improves the aerodynamics of the turbine vane (34).

Abstract: Within components such as high pressure turbine blades and aerofoils in a gas turbine engine it is important to provide cooling such that these components remain within acceptable operational parameters. Typically, film cooling as well as convective cooling is utilized. Film cooling requires holes from a feed passage from which the coolant is presented upon an external surface to develop the film. The holes themselves can create cooling through convective cooling effects. In order to maximize the convective cooling effect holes are created which have an indirect path about a direct line between an inlet and an outlet for the hole. By creating an indirect path in the form of a helix or spiral which in turn may have a variable cross sectional area from the inlet to the outlet control of coolant flow can be achieved.

Abstract: Within aerofoils, and in particular nozzle guide vane aerofoils in gas turbine engines problems can occur with regard to coolant flows from respective inlets at opposite ends of a cavity within the aerofoil. The cavity generally defines a hollow core and unless care is taken coolant flow can pass directly across the internal cavity. Previously baffle plates were inserted within the cavity to prevent such direct jetting across the cavity. Such baffle plates are subject to additional costs as well as potential unreliability problems. Baffles formed integrally with a wall within the aerofoil allow more reliability with regard to positioning as well as consistency of performance. The baffles can be perpendicular, upward or downwardly orientated or have a compound angle.

Abstract: Within aerofoils, and in particular nozzle guide vane aerofoils in gas turbine engines problems can occur with regard to coolant flows from respective inlets at opposite ends of a cavity within the aerofoil. The cavity generally defines a hollow core and unless care is taken coolant flow can pass directly across the internal cavity. Previously baffle plates were inserted within the cavity to prevent such direct jetting across the cavity. Such baffle plates are subject to additional costs as well as potential unreliability problems. Baffles formed integrally with a wall within the aerofoil allow more reliability with regard to positioning as well as consistency of performance. The baffles can be perpendicular, upward or downwardly orientated or have a compound angle.

Abstract: Within components such as high pressure turbine blades and aerofoils in a gas turbine engine it is important to provide cooling such that these components remain within acceptable operational parameters. Typically, film cooling as well as convective cooling is utilised. Film cooling requires holes from a feed passage from which the coolant is presented upon an external surface to develop the film. The holes themselves can create cooling through convective cooling effects. In order to maximise the convective cooling effect holes are created which have an indirect path about a direct line between an inlet and an outlet for the hole. By creating an indirect path in the form of a helix or spiral which in turn may have a variable cross sectional area from the inlet to the outlet control of coolant flow can be achieved.

Abstract: A cooling arrangement 21 for use within a gas turbine engine comprises a first shroud or platform 26 incorporating coolant passages 25 and a second shroud or platform 28. Generally, each platform or shroud 26, 28 will incorporate a pressure portion and a suction portion, with the pressure portion incorporating the coolant passages 25 through which the coolant flow 27 becomes incident on a surface 40 of the suction portion of the second shroud 28. The surface 40 is inclined or tapered towards the passage 25, such that there is limited direct impingement upon a front edge 39 of the surface 40. The coolant flow 27 thereby remains adjacent to the surface 40 for a longer period and so enhances cooling efficiency.

Abstract: A cooling and sealing arrangement in a gas turbine engine comprises a seal assembly extending between nozzle guide vanes and combustion chamber discharge nozzles. The nozzle guide vanes have platforms having extensions which in one embodiment overlap the downstream ends of the discharge nozzles. Cooling air is supplied to the upstream edge portions of the platforms through holes in flanges on the discharge nozzles. The seal assembly defines a chamber adjacent the platform extensions which is supplied with cooling air through metering holes in the seal assembly which provides part of a boundary for a cooling air chamber, the boundary of the chamber having cooling air holes for metering cooling air into the chamber to cool the upstream portion of the nozzle guide vane outer platforms. Cooling air exits the chamber through holes in the platform extensions to film cool them.

Abstract: A turbine nozzle assembly includes an annular array of nozzle guide vanes located downstream of a combustor discharge casing. Each nozzle guide vane includes an aerofoil portion which is cast integrally with a radially inner platform and a radially outer platform. The radially outer platform of each nozzle guide vane has an extension to provide a smooth transition of the gases from the combustor discharge casing to the nozzle guide vanes. Two rows of cooling holes are provided in the extension to film cool the inner surface of the platform. A method is described to calculate the diameter of each of the cooling holes so that a uniform flow of cooling air passes over the inner surface of the each platform.

Abstract: A cooling and sealing arrangement in a gas turbine engine comprises a seal assembly (42,30) extending between nozzle guide vanes (20) and combustion chamber discharge nozzles (22). The nozzle guide vanes (20) have platforms (34,36) having extensions (34A,36A) which in one embodiment overlap the downstream ends of the discharge nozzles (22). Cooling air is supplied to the upstream edge portions of the platforms through holes (26,28C) in flanges (24,28) on the discharge nozzles (22). The seal assembly (42,30) defines a chamber (58,70) adjacent the platform extensions which is supplied with cooling air through metering holes (56,68) in the seal assembly. Cooling air exits the chamber (58,70) through holes (34B,36B) in the platform extensions to film cool them.