Abstract: A method of detecting a deviation angle in a single-crystal metal structure is disclosed. The single-crystal metal structure has a crystallographic orientation, a length, a first side, a second side, and a first axis extending through the structure. The method comprises determining the length of the single-crystal metal structure along the first axis, transmitting a signal through the single-crystal metal structure from the first side, the signal oriented to propagate along the first axis, receiving the signal, determining a time-of-flight for the signal to traverse the length from the first side to the second side, determining a speed of the signal based on the time-of-flight and the length, and comparing the speed of the signal to a reference speed to detect the deviation angle.

Abstract: In accordance with an exemplary embodiment, a turbine stator component includes a first endwall; a second endwall; a first stator airfoil coupled between the first and second endwalls; and a second stator airfoil adjacent to the first airfoil and coupled between the first and second endwalls. The first stator airfoil has first crystallographic primary and secondary orientations. The second stator airfoil has second crystallographic primary and secondary orientations, the first crystallographic primary and secondary orientations being different from the second crystallographic primary and secondary orientations.

Abstract: Turbine blades and methods of forming modified turbine blades and turbine rotors for use in an engine are provided. In an embodiment, by way of example only, a turbine blade includes a platform and an airfoil. The platform includes a surface configured to define a portion of a flowpath, and the surface includes an initial contour configured to plastically deform into an intended final contour after an initial exposure of the blade to an operation of the engine. The airfoil extends from the platform.

Abstract: A turbine blade assembly is provided. The turbine blade assembly comprises a turbine blade comprising a cavity, and a blade platform supporting the turbine blade, the cavity extending into the blade platform. The blade platform comprises an upper surface adjacent the turbine blade and a lower surface comprising a first rib, the cavity extending into the first rib, the first rib coupled to the lower surface, tapering as it extends away from the turbine blade, and comprising a first port extending from the cavity to the upper surface.

Abstract: A method of detecting a deviation angle in a single-crystal metal structure is disclosed. The single-crystal metal structure has a crystallographic orientation, a length, a first side, a second side, and a first axis extending through the structure. The method comprises determining the length of the single-crystal metal structure along the first axis, transmitting a signal through the single-crystal metal structure from the first side, the signal oriented to propagate along the first axis, receiving the signal, determining a time-of-flight for the signal to traverse the length from the first side to the second side, determining a speed of the signal based on the time-of-flight and the length, and comparing the speed of the signal to a reference speed to detect the deviation angle.

Abstract: Turbine nozzles and methods of manufacturing the turbine nozzles are provided. In an embodiment, by way of example only, a turbine nozzle includes a first ring, a vane, and a first joint. The first ring comprises a single unitary component and having a first opening and including a first metal alloy. The vane includes a first end disposed in the first opening and includes a second metal alloy. The first joint is formed in the first opening between the first ring and the vane and includes a first braze layer and an oxide layer. The first braze layer is disposed adjacent to the oxide layer, and the first braze layer and the oxide layer are disposed between the first ring and the vane.

Abstract: Spacers and turbines are provided. In an embodiment, and by way of example only, a spacer includes a strip, a first retention flange, and a second retention flange. The strip has a first edge, a second edge, and an impingement surface, and the impingement surface extends axially along the strip between the first edge and the second edge and is substantially flat. The first retention flange is recessed relative to the impingement surface and extends away from the first edge of the strip.

Abstract: A gas turbine engine assembly has combustion gases flowing through a gas flow path. The gas turbine engine assembly includes a stator assembly comprising a stator vane that extends into the gas flow path; and a turbine rotor assembly downstream of the stator assembly and comprising a turbine platform and a turbine rotor blade extending from the turbine platform into the mainstream combustion gases flow path. The turbine rotor blade includes a pressure side and a suction side opposing the pressure side that extend from a leading edge to a trailing edge. The combustion gases form horseshoe vortices at a formation area adjacent the leading edge of the turbine rotor blade, and the turbine rotor assembly further includes a first set of holes in the turbine platform for directing first jets into the formation area of the horseshoe vortices.

Abstract: A gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path configured to receive mainstream hot gas flow. The assembly further includes a stator assembly including a stator vane that extends into the mainstream hot gas flow path and a turbine rotor assembly downstream of the stator assembly that includes a turbine disk and a turbine blade extending from the turbine disk into the mainstream hot gas flow path. The stator assembly and turbine assembly define a turbine disk cavity, and the turbine disk cavity includes a recirculation cavity configured to recirculate gas ingested from the mainstream hot gas flow path back into the mainstream hot gas flow path.

Abstract: A gas turbine engine includes a housing with a duct wall that defines a generally annular and axially elongated hot gas flow path for passage of combustion gas. The engine further includes a contoured shroud mounted within the internal engine cavity and defining a hot gas recirculation pocket for receiving hot gas ingested from the hot gas flow path through the annular space and for recirculating the ingested hot gas back through the annular space to the hot gas flow path. The contoured shroud includes a base wall extending radially inwardly from the duct wall, an inboard wall extending from the base wall in an axial direction toward the rotor, and an end wall extending from the inboard wall in a radially outward direction. The end wall terminates in a circumferentially extending free edge disposed in proximity to the annular space. The end wall defines an opening.

Abstract: An air-cooled turbine blade is provided having an airfoil shape defined by a convex suction side wall, a concave pressure side wall, a leading edge, a trailing edge, a root and a tip, the walls and the tip each including an interior surface that defines an interior with the root, the trailing edge including a plurality of slots formed thereon. The blade includes a suction side flow circuit formed therein comprising a forward and an aft flow circuit. The blade also includes a tip flow circuit extending along the tip interior surface to at least one of the trailing edge slots and including a first and a second opening, the first opening in flow communication with the suction side forward flow circuit outlet, and the second opening in flow communication with a cross-over hole of the suction side aft flow circuit. Methods of manufacturing the blade are also provided.

Abstract: A core assembly including two cores is used to manufacture a blade. The first core has an outer surface shaped to complement the tip wall bottom surface. The second core has a tip surface, a side surface, and a protrusion or a depression. The tip surface is shaped to complement at least a portion of the tip wall top surface and is configured to be disposed proximate the first core. The side surface is shaped to complement at least a portion of the side wall, and the protrusion extends from the second core side surface to contact at least a portion of the ceramic mold inner surface. In embodiments employing a depression, the depression is formed in the side surface.

Abstract: A diffuser particle separator may be integrated into a gas turbine engine to remove corrosive dust and salt particles from the engine's core air flow. The air flow may pass over a series of particle accumulator entrance orifices, trapping particles in a particle accumulator while allowing the air flow to continue unimpeded. Since dust deposits may become molten at high temperatures, removal of dust from the core and secondary airflow may be critical for long-life superalloy and ceramic components, particularly those with small diameter air-cooling holes and thermal barrier coatings.

Abstract: Methods are provided for modifying an airfoil internal cooling circuit that include a flow path configured to direct air through the airfoil in a direction and the airfoil having a leading edge, a trailing edge, and a first and a second wall therebetween, each wall having an inner and an outer surface, the inner surfaces defining a cavity and having features forming at least a portion of the internal cooling circuit. The methods may include the steps of forming a pilot hole through the airfoil first and second walls at a predetermined location, forming an insert hole based on the predetermined location, the insert hole enveloping the pilot hole and configured to receive at least a portion of an insert configured to modify the internal cooling circuit flow path, placing the insert into the insert hole, and bonding the insert to the airfoil first and second walls.

Abstract: A core assembly including two cores is used to manufacture a blade. The first core has an outer surface shaped to complement the tip wall bottom surface. The second core has a tip surface, a side surface, and a protrusion or a depression. The tip surface is shaped to complement at least a portion of the tip wall top surface and is configured to be disposed proximate the first core. The side surface is shaped to complement at least a portion of the side wall, and the protrusion extends from the second core side surface to contact at least a portion of the ceramic mold inner surface. In embodiments employing a depression, the depression is formed in the side surface.

Abstract: An air-cooled turbine blade is provided having an airfoil shape defined by a convex suction side wall, a concave pressure side wall, a leading edge, a trailing edge, a root and a tip, the walls and the tip each including an interior surface that defines an interior with the root, the trailing edge including a plurality of slots formed thereon. The blade includes a suction side flow circuit formed therein comprising a forward and an aft flow circuit. The blade also includes a tip flow circuit extending along the tip interior surface to at least one of the trailing edge slots and including a first and a second opening, the first opening in flow communication with the suction side forward flow circuit outlet, and the second opening in flow communication with a cross-over hole of the suction side aft flow circuit. Methods of manufacturing the blade are also provided.

Abstract: A method of modifying a turbine nozzle area comprises depositing a thermal barrier coating (TBC) on the nozzle endwalls to provide a minimum nozzle area, evaluating an airflow through the nozzle, and machining the TBC to increase the nozzle area. Adjacent segment area variation may be minimized, improving engine reliability by reducing the aerodynamic excitation to the down stream blade.

Abstract: A gas turbine engine is configured to use relatively cool, low pressure air discharged from a low pressure compressor to supply buffer air to lubrication sump seals. The engine is further configured such that the lubrication sump is thermally layered by isolating relatively hot, high pressure compressor air from the sump by utilizing a warm vent mixing cavity, which is located radially between of the hot high pressure compressor air and the cool buffer air, which is located in a buffer cavity between the vent cavity and the sump.

Abstract: A turbine shroud assembly includes forward and aft hangers, an axisymmetric plenum assembly, ceramic shroud segments, ceramic spacers, and forward and aft rope seals. The plenum assembly supplies impingement cooling to the shroud and the hangers. The impingement cooling to the forward and aft hangers is controlled independently to improve blade tip clearance. The rope seals are radially inward from the hangers and reduce cooling flow leakage. The turbine shroud assembly can operate in a higher temperature environment using less cooling flow than the prior art.

Abstract: A robust multiple-walled, multi-pass, high cooling effectiveness cooled turbine vane or blade designed for ease of manufacturability, minimizes cooling flows on highly loaded turbine rotors. The vane or blade design allows the turbine inlet temperature to increase over current technology levels while simultaneously reducing turbine cooling to low levels. A multi-wall cooling system is described, which meets the inherent conflict to maximize the flow area of the cooling passages while retaining the required section thickness to meet the structural requirements. Independent cooling circuits for the vane or blade's pressure and suction surfaces allow the cooling of the airfoil surfaces to be tailored to specific heat load distributions (that is, the pressure surface circuit is an independent forward flowing serpentine while the suction surface is an independent rearward flowing serpentine). The cooling air for the independent circuits is supplied through separate passages at the base of the vane or blade.