Abstract: An apparatus for cooling turbine blades in a turbine engine including a direct transfer axial tangential onboard injector (TOBI) for a turbine rotor and a self-supporting seal plate disposed on a rotating disk for the turbine engine. The TOBI includes a plurality of openings emanating a flow of cooling air. The self-supporting seal plate comprises a plurality of shaped cooling holes in fluid communication with the flow of cooling air emanating from the TOBI. The rotating disk includes a plurality of turbine blade slots formed therein. The plurality of cooling holes are in fluid communication with the plurality of turbine blade slots for directing the flow of cooling air to provide cooling to the plurality of turbine blades. The plurality of openings, the plurality of cooling holes and the plurality of turbine blade slots are in axial alignment and optimized to minimize radial and hoop stresses.

Abstract: A seal plate for a turbine engine is provided. The turbine engine has a central axis. The seal plate comprises a forward face extending radially from the central axis. The forward face comprises a flat portion extending circumferentially around the central axis, and a front flow discourager coupled to the flat portion and extending outward from the flat portion.

Abstract: A coupling assembly for a turbine shroud is provided. The coupling assembly comprises a rotatable positioning block having a first surface, and a biasing spring having a second surface, the second surface generally facing the first surface, and the biasing spring adapted to exert a force toward the positioning block when compressed.

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: 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: 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: 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: 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, each wall including an interior surface that defines an interior with the edges, root and tip. The blade includes a plurality of independent cooling circuit flow paths within the blade interior and a roughened surface. The roughened surface is formed on an interior surface of at least one of the convex suction side wall and the concave pressure side wall, the roughened surface comprising a plurality of cylindrical depressions. Each depression includes a cylindrical wall coupled to a bottom wall by a chamfered edge formed therebetween, and the roughened surface defines at least a portion of one of the flow paths of the plurality of independent cooling circuit flow paths. Methods for forming the blade are also provided.

Abstract: A compressor scroll housing for use in conjunction with turbo-machinery, particularly applicable in aircraft. The scroll housing can include a plurality of scroll vanes arrayed around the scroll housing. Scroll vanes, integrally formed with the scroll housing, carry stress load on the scroll housing, including the load from fluid pressure within the scroll and carcass loading from the engine. The plurality of scroll vanes adapted for guiding flow of fluid from an inlet to an outlet while supporting the scroll housing. Chord length and cross sectional area of each scroll vane can be sized to maintain an equal stress in all scroll vanes. A method of making the scroll housing for use with an impeller connected to an engine is disclosed, as well as a method of operating turbo-machinery including supporting a load on the scroll housing with scroll vanes while maintaining an equal stress on each scroll vane.

Abstract: A turbine engine compressor design utilizing multiple component integration, thereby reducing the number of required engine components. In conventional compressor designs, a multiple component system makes it difficult to predict the structural behaviors due to thermal and mechanical loading during transient conditions. The compressor design of the present invention has three main parts: a forward bearing housing, a bell-mouth (heat shield) and a coupled impeller shroud/diffuser. Such a design achieves the design objectives of the present invention, including reducing weight, reducing cost, minimizing tolerance build up and improving aerodynamic performance by utilizing multiple component integration for multiple modes of engine operation.