Abstract: A composite core die includes a reusable core die and a disposable core die; where the disposable core die is in physical communication with the reusable core die; and further where surfaces of communication between the disposable core die and the reusable core die serve as barriers to prevent the leakage of a slurry that is disposed in the composite core die.

Abstract: Hybrid components containing a ceramic material, in which detailed features of the components are formed of materials other than ceramic materials, yet result in a robust mechanical attachment of the ceramic and non-ceramic portions of the components. The components includes a first subcomponent formed of a ceramic-based material and a second subcomponent formed of a metallic material. The first subcomponent has a nub and the second subcomponent is separately formed and attached to the first subcomponent by casting the metallic material around the nub of the first subcomponent. The second subcomponent is attached to the first subcomponent by a compression fit and encapsulation of the second subcomponent on the nub of the first subcomponent. The nub has a compliant coating system that provides thermal expansion compliance between the metallic material of the second subcomponent and the ceramic-based material of the first subcomponent.

Abstract: Hybrid turbine airfoil components containing a ceramic material, in which detailed features of the components are formed of materials other than ceramic materials. The components include a first component formed of a ceramic-based material and a second component formed of a metallic material. The first component comprises an airfoil portion and a nub, and the second component is separately formed and attached to the first component by casting the metallic material around the nub of the first component. The second component includes a platform portion between the airfoil portion and the nub of the first component and a dovetail portion on the nub of the first component. Each of the platform and dovetail portions has at least one off-axis geometric feature that results in the second component having a more complex geometry than the first component.

Abstract: Processes for producing a component containing a ceramic-based material and having detailed features formed from materials other than ceramic materials. Such a process entails producing the component to include a first subcomponent and at least a second subcomponent having at least one off-axis geometric feature that results in the second subcomponent having a more complex geometry than the first subcomponent. The first subcomponent is formed of a ceramic-based material, and the second subcomponent and its off-axis geometric feature are separately formed of a metallic material and attached to the first subcomponent to yield a robust mechanical attachment. The component may be, for example, a gas turbine airfoil component.

Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface, and coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment.

Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip to facilitate film cooling the shroud segment.

Abstract: A method of assembling a gas turbine engine includes coupling a turbine shroud assembly within the gas turbine engine. The turbine shroud assembly includes a shroud segment having a leading edge defining a forward face and a radial inner surface. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face. A plurality of recuperated cooling openings are defined through the leading edge at an oblique inlet angle with respect to a centerline of the gap and between the forward face and the radial inner surface to direct cooling fluid through the leading edge to facilitate preferential cooling of the leading edge.

Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip. Post impingement cooling air is directed at the forward face to facilitate forming a film cooling layer on the shroud segment. The film cooling layer is shielded from combustion gases flowing through the gas turbine engine.

Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, and directing the cooling air in the gap through at least one cooling hole extending between the leading edge and the inner surface.

Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface.

Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, circumferentially mixing the cooling air in a plenum defined within the leading edge to substantially uniformly distribute the cooling air throughout the gap, and directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface.

Abstract: Disclosed herein is a method comprising injecting into a thin wall disposable core die a slurry having a viscosity of about 1 to about 1,000 Pascal-seconds at room temperature when tested at a shear rate of up to 70 seconds?1 and a flow index of less than 0.6 at a pressure of up to about 7 kilograms-force per square centimeter; wherein the thin wall disposable core die has an average wall thickness of about 1.5 to about 10 millimeters; curing the slurry to form a cured ceramic core; removing the thin wall disposable core die from the cured ceramic core; and firing the cured ceramic core to form a solidified ceramic core.

Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, circumferentially mixing the cooling air in a plenum defined within the leading edge to substantially uniformly distribute the cooling air throughout the gap, and directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface.

Abstract: Disclosed herein is a composite core die comprising a reusable core die; and a disposable core die; wherein the disposable core die is in physical communication with the reusable core die; and further wherein surfaces of communication between the disposable core die and the reusable core die serve as barriers to prevent the leakage of a slurry that is disposed in the composite core die.

Abstract: Disclosed herein is a method comprising injecting into a thin wall disposable core die a slurry having a viscosity of about 1 to about 1,000 Pascal-seconds at room temperature when tested at a shear rate of up to 70 seconds?1 and a flow index of less than 0.6 at a pressure of up to about 7 kilograms-force per square centimeter; wherein the thin wall disposable core die has an average wall thickness of about 1.5 to about 10 millimeters; curing the slurry to form a cured ceramic core; removing the thin wall disposable core die from the cured ceramic core; and firing the cured ceramic core to form a solidified ceramic core.

Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface.

Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface, and coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment.

Abstract: A method of assembling a gas turbine engine includes coupling a turbine shroud assembly within the gas turbine engine. The turbine shroud assembly includes a shroud segment having a leading edge defining a forward face and a radial inner surface. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face. A plurality of recuperated cooling openings are defined through the leading edge at an oblique inlet angle with respect to a centerline of the gap and between the forward face and the radial inner surface to direct cooling fluid through the leading edge to facilitate preferential cooling of the leading edge.

Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, and directing the cooling air in the gap through at least one cooling hole extending between the leading edge and the inner surface.

Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip to facilitate film cooling the shroud segment.