Abstract: Modular assemblies for gas turbine engines such as modular vane assemblies and methods of manufacturing the same. The modular assembly includes a first modular component such as a vane platform having a first mating pocket, and a second modular such as an airfoil. The second modular component includes circumferentially extending first and second surfaces at first and second distal ends thereof, respectively, with the first surface being received within the first pocket when the modular assembly is in the assembled state. The second modular component also includes a coating pocket extending from the first surface to the second surface. The coating pocket is recessed towards an interior of the second modular component with respect to first surface and the second surface, and a thermal barrier coating is included within the coating pocket and not included on the first surface or the surface.

Abstract: Modular assemblies for gas turbine engines such as modular vane assemblies and methods of manufacturing the same. The modular assembly includes a first modular component such as a vane platform having a first mating pocket, and a second modular such as an airfoil. The second modular component includes circumferentially extending first and second surfaces at first and second distal ends thereof, respectively, with the first surface being received within the first pocket when the modular assembly is in the assembled state. The second modular component also includes a coating pocket extending from the first surface to the second surface. The coating pocket is recessed towards an interior of the second modular component with respect to first surface and the second surface, and a thermal barrier coating is included within the coating pocket and not included on the first surface or the surface.

Abstract: Airfoils, gas turbine engine assemblies including such airfoils, and methods of manufacturing the same. The airfoil includes multiple cooling channels proximate a leading edge of the airfoil, each of the cooling channels including an inlet provided on one an inner surface of the airfoil in one of the pressure side wall and the suction side wall, and an outlet provided on an outer surface of the airfoil in the other of the pressure side wall and the suction side wall. The cooling channels are arranged in a staggered configuration such there is an alternate pattern of cooling fluid flow provided at the leading edge of the airfoil, near the airfoil's stagnation point.

Abstract: An airfoil and turbine vanes and vane assemblies incorporating the same. The airfoil has an uncoated profile substantially in accordance with Cartesian coordinate values of X, Y, and Z as set forth in Table 1, carried to four decimal places. The Z values refer to a percentage of the radial span of the airfoil measured radially from a radially outwardly facing surface of the inner platform. The turbine vane includes an inner platform, an outer platform, and an airfoil such as the one discussed above extending radially outward from the inner platform toward the outer platform. And the vane assembly includes an inner platform, an outer platform, and two or more first stage vanes extending from the inner platform to the outer platform. Each of the two or more first stage vanes include an airfoil as discussed above.

Abstract: An airfoil and turbine vanes and vane assemblies incorporating the same. The airfoil has an uncoated profile substantially in accordance with Cartesian coordinate values of X, Y, and Z as set forth in Table 1, carried to four decimal places. The Z values refer to a percentage of the radial span of the airfoil measured radially from a radially outwardly facing surface of the inner platform. The turbine vane includes an inner platform, an outer platform, and an airfoil such as the one discussed above extending radially outward from the inner platform toward the outer platform. And the vane assembly includes an inner platform, an outer platform, and two or more first stage vanes extending from the inner platform to the outer platform. Each of the two or more first stage vanes include an airfoil as discussed above.

Abstract: A method for manufacturing and heat treating a component formed from an alloy or superalloy. The component may be part of a gas turbine engine, such as an airfoil of a gas turbine engine turbine, or a combustor fuel nozzle, that includes internal cooling passages. At least part of the component is formed from the alloy or superalloy using an additive manufacturing process. The method includes determining an incipient melting point range and a recrystallization temperature of the alloy or superalloy. The component is then heated to at least the recrystallization temperature and within the incipient melting point range for a predetermined time, and ultimately cooled.

Abstract: Airfoils, gas turbine engine assemblies including such airfoils, and methods of manufacturing the same. The airfoil includes multiple cooling channels proximate a leading edge of the airfoil, each of the cooling channels including an inlet provided on one an inner surface of the airfoil in one of the pressure side wall and the suction side wall, and an outlet provided on an outer surface of the airfoil in the other of the pressure side wall and the suction side wall. The cooling channels are arranged in a staggered configuration such there is an alternate pattern of cooling fluid flow provided at the leading edge of the airfoil, near the airfoil's stagnation point.

Abstract: A method for manufacturing and heat treating a component formed from an alloy or superalloy. The component may be part of a gas turbine engine, such as an airfoil of a gas turbine engine turbine, or a combustor fuel nozzle, that includes internal cooling passages. At least part of the component is formed from the alloy or superalloy using an additive manufacturing process. The method includes determining an incipient melting point range and a recrystallization temperature of the alloy or superalloy. The component is then heated to at least the recrystallization temperature and within the incipient melting point range for a predetermined time, and ultimately cooled.

Abstract: A system and method of extending the useable life of a gas turbine blade is disclosed in which the gas turbine blade includes an undercut configuration designed to relieve mechanical and thermal stress imparted into the pedestal region of the airfoil trailing edge. The embodiments of the present invention include turbine blade configurations having different trailing edge undercut configurations as well as additional cooling supplied to the internal passages of the trailing edge region of the turbine blade.

Abstract: A system and method of extending the useable life of a gas turbine blade is disclosed in which the gas turbine blade includes an undercut configuration designed to relieve mechanical and thermal stress imparted into the pedestal region of the airfoil trailing edge. The embodiments of the present invention include turbine blade configurations having different trailing edge undercut configurations as well as additional cooling supplied to the internal passages of the trailing edge region of the turbine blade.