Abstract: A one-piece deflector-flare cone assembly for a gas turbine engine combustor that facilitates extending a useful life of the combustor in a cost-effective and reliable manner is described. The combustor includes an air swirler, and the one-piece assembly includes a deflector portion and a flare cone portion. The deflector portion includes an integral opening that extends circumferentially through the deflector portion for receiving cooling fluid therein. Because the deflector-flare cone assembly receives pre-loaded braze tape and braze rope, a single braze operation secures the deflector-flare cone assembly within the combustor and to a combustor air swirler.

Abstract: A turbofan engine includes a pulse detonation system to create a temperature rise and a pressure rise within the engine to generate thrust from the engine. The system includes a pulse detonation augmentor including a shock tube sub-system. The shock tube sub-system includes a plurality of shock tubes which mix air and fuel introduced to the pulse detonation augmentor and detonate the mixture. The detonation creates hot combustion gases which are directed from the engine to produce thrust for the engine. Alternatively, the system includes a pulse detonation augmentation system that replaces a core engine of a turbo-fan engine.

Abstract: A combustor liner has a stepped combustor liner surface defining a combustion zone and an overhang portion forming an air cooling slot. A layer of thermal barrier material is applied to the combustor liner such that at least one portion of the combustor liner receives a layer of thermal barrier material with a thickness greater than 0.01 inches. Thus, the combustor liner absorbs less heat, and the combustor may operate at higher temperatures. As a result, low cycle fatigue and thermal creep are reduced within the combustor and the life cycle for the combustor is extended.

Abstract: An augmenter includes at least one igniter which provides a uniform ignition within the augmenter. The igniter is secured to an igniter housing with a self-locking nut and each igniter includes a biasing mechanism which biases the igniter against the augmenter. Biasing the igniter against the augmenter ensures that each igniter tip is maintained at a proper immersion depth within the augmenter during gas turbine engine operation. Additionally, the igniter includes an alignment support boss to ensure that the igniter tip is properly aligned within the augmenter as the igniter is attached to the augmenter.

Abstract: A stator includes a row of airfoils extending inwardly from a casing. The airfoils are spaced apart circumferentially from each other at a spacing varying in turn around the perimeter of the casing for reducing vibratory excitation of a downstream rotor stage.

Abstract: An air bleed assembly for extracting air from a flowpath in a gas turbine engine includes a casing for surrounding a row of circumferentially spaced apart rotor blades and defining at least in part the flowpath for receiving air compressed by the rotor blades. The casing includes a bleed port disposed downstream of a row of the blades for receiving a portion of the compressed air as bleed airflow and a bleed duct extending away from the bleed port. The bleed duct has an annular slot in the casing. The annular slot has annular slot leading and trailing edges and an annular bifurcated splitter disposed along at least a portion of the annular slot trailing edge. The bifurcated splitter has an annular leading edge forebody located upstream of and separated from an annular splitter wall by an annular return channel. In one embodiment of the invention, the rotor blades are fan blades and the annular bleed space is in the flowpath in a fan section of the engine.

Abstract: A nozzle vane includes a row of trailing edge apertures and cooperating inboard slots joined in flow communication with a mid-chord cavity. An outboard slot is spaced outwardly from a respective last one of the inboard slots, and outboard of a respective last one of the trailing edge apertures. The outboard slot extends behind a braze joint between the vane and a supporting band and is effective for backside cooling thereof.

Abstract: An article assembly comprising a plurality of members includes a multi-layered brazed joint provided in the assembly between a first member that can be affected adversely by thermal exposure at a limiting and a second member requiring exposure, such as for brazing, at least at the limiting temperature. The assembly is provided with a multi-layered brazed joint directly or indirectly between such members by brazing to the second member a second brazing alloy that can be brazed at a temperature required to braze the second member, providing a preform. The assembly is brazed to include the first member by brazing members of the assembly together with a first brazing alloy, brazing at a temperature less than the limiting temperature, brazed at the second brazing alloy.

Abstract: Disengagement of the C-clip in turbine shroud assemblies is prevented by providing a C-clip retainer. The shroud assembly includes a shroud support having a hook and at least one shroud having a mounting flange. A C-clip overlaps the hook and the mounting flange to clamp the shroud to the shroud support, and the retainer is secured to the shroud support and located so as to engage the C-clip. The retainer engages the C-clip in such a manner so as to limit aft axial movement of the C-clip, thereby eliminating C-clip back-off.

Abstract: A turbine for a gas turbine engine including a turbine nozzle assembly that facilitates reducing an operating temperature of rotor blades in a cost-effective and reliable manner is described. Each rotor blade includes a tip that rotates in close proximity to a shroud that extends circumferentially around the rotor assembly. The turbine nozzle assembly includes a plurality of turbine vane segments that channel combustion gases to downstream rotor blades. Each turbine vane segment extends radially outward from an inner platform and includes a tip, a root, and a body that extends therebetween. The turbine vane segment tip is formed integrally with an outer band that mounts the vane segments within the gas turbine engine. The outer band is in flow communication with a cooling fluid source, and includes at least one opening.

Abstract: A turbine nozzle includes vanes joined to outer and inner bands. Each vane includes opposite sidewalls extending between leading and trailing edges, with an internal bridge extending between the sidewalls to effect forward and aft cavities. Side ribs extend along the sidewalls and project inwardly into the cavities. Each of the side ribs is truncated adjacent the bridge to uncouple the ribs therefrom and increase nozzle life.

Abstract: A combustor for a gas turbine engine operates with high combustion efficiency and low carbon monoxide, nitrous oxide, and smoke emissions during low, intermediate, and high engine power operations is described. The combustor includes a mixer assembly including a pilot mixer, a main mixer, and a mid-power and cruise mixer. The pilot mixer includes a pilot fuel injector, at least one swirler, and an air splitter. The main mixer extends circumferentially around the pilot mixer. The mid-power and cruise mixer extends between the main and pilot mixers and includes a plurality of fuel injection ports which inject fuel radially inwardly to facilitate radial and circumferential fuel-air mixing to provide a substantially uniform fuel and air distribution for combustion.

Abstract: An aircraft gas turbine axisymmetric vectoring nozzle has an interface ring centered about a nozzle, a vectoring ring disposed radially inwardly of and apart from the interface ring, and a bearing radially disposed between the vectoring ring and the interface ring. The bearing may be a sliding bearing having a sliding interface between the vectoring ring and the interface ring and the sliding interface is spherical. The bearing may be constructed of sliding bearing segments having sliding interfaces between the vectoring ring and the interface ring and the sliding interfaces are spherical. Each of the bearing segments includes an outer sliding element attached to the interface ring, an inner sliding element attached to the vectoring ring, and spherically curved outer and inner sliding surfaces on the outer and inner sliding elements respectively wherein the spherically curved outer and inner sliding surfaces define the sliding interfaces between the vectoring ring and the interface ring.

Abstract: A heat shield for protecting a fuel injection tube of a turbine engine afterburner. The heat shield includes a base adapted for attachment to a duct member, the base having a first mounting surface, a first fastener hole, and a recess in the base. A housing encloses the fuel injection tube. The housing has an internal channel for receiving the fuel injection tube, a second mounting surface, a second fastener hole, and a tab protruding from the housing. The tab is sized and shaped for being received in the recess of the base. The housing may be secured to the base at a mounting position wherein the tab is received in the recess, the mounting surfaces are engaged, and the second fastener hole is positioned in registration with the first fastener hole whereby a fastener may be inserted to secure the housing to the base.

Abstract: A tool for removing a shim from a dovetail of a fan blade for a turbine engine rotor includes a clip and a handle. The clip includes a release edge configured to engage a flared lip of a dovetail shim. The handle extends from the clip and contacts the exterior surface of the shim. When the handle and the clip are rotated about a longitudinal axis of the handle, the release edge separates the flared lip of the shim from the dovetail and the handle applies a downward force to easily remove the shim from the dovetail.

Abstract: A gas turbine engine assembly including a gas turbine engine mounted within a module and including an inlet in flow communication with a module inlet area and a module exhaust area is described. The turbine engine also includes an exhaust, and is mounted within an engine area between the module inlet and exhaust areas. More specifically, the gas turbine engine is mounted such that the engine exhaust is in flow communication with the module exhaust area. Because the module also includes a secondary supply duct, the gas turbine engine is in flow communication with the module inlet and exhaust areas.

Abstract: Disengagement of C-clips in turbine shroud assemblies is prevented by providing each C-clip with at least one anti-rocking pad. The anti-rocking pad is disposed on a radially inner surface of the C-clip so as to make snug contact with a portion of the shroud adjacent to the C-clip. The snug contact limits rocking motion of the C-clip, thereby preventing C-clip disengagernent.

Abstract: A gas turbine engine includes a lubrication system that supplies oil through a plurality of circumferential grooves and radial grooves to a bearing assembly. The grooves extend within an inner surface of a first rotor shaft sized to fit in an interference fit around a main rotor shaft. The circumferential grooves are spaced circumferentially around the inner surface of the first rotor shaft and are substantially perpendicular to the radial grooves. The radial grooves are in flow communication with a plurality of scoops extending between an outer and inner surface of the first rotor shaft, and with the bearing assembly.

Abstract: A turbine airfoil includes at least one spar arrangement having a length less than an associated turbine airfoil length. During operation, the turbine airfoil has an outer skin surface which operates at a substantially higher temperature than that of an internal supporting parted spar arrangement. The parted spar arrangement permits the turbine airfoil outer skin surface to thermally expand between spar arrangements, thus preventing self-constraining thermal stresses from forming within the spar arrangement or the airfoil skin surfaces.

Abstract: A gas turbine engine rotor assembly includes a plurality of aerodynamic devices to direct airflow radially inward. The gas turbine engine rotor assembly includes a rotor shaft that includes a plurality of openings. The aerodynamic devices include a pair of vane segments and a pair of sidewalls. A contoured outer surface includes an opening and permits the aerodynamic device to be positioned against an inner surface of the rotor shaft, and a flange ring defines a pocket. The aerodynamic device fits within the pocket to concentrically align the openings.