Abstract: A turbine shroud includes a panel having inner and outer surfaces extending between forward and aft opposite ends. The panel includes forward and aft hooks for supporting the panel radially atop a row of turbine rotor blades. The panel includes a plurality of recesses in the inner surface thereof which face tips of the blades. The recesses extend only in part into the panel for reducing surface area exposed to the blade tips.

Abstract: A one-piece cowl for use in assembled relationship with a combustor of a gas turbine engine, the cowl being of a generally annular configuration defining a central cowl axis and being axially elongated and aerodynamically contoured relative to the central cowl axis. The one-piece cowl includes an inner annular portion, an outer annular portion, and a plurality of circumferentially spaced radial ligaments connecting the inner and outer annular portions so as to form a corresponding number of openings through a middle portion of the cowl. Each radial ligament has a designated area which enables the radial ligaments to sustain the stress levels imposed on the cowl to prevent high cycle fatigue. Each opening also has a designated area which enables insertion of fuel nozzles therein without adversely affecting air flow therethrough. The designated area for the radial ligaments is a function of the designated area for the openings and a ratio of such areas falls within a specified range.

Abstract: A fuel nozzle including a nozzle stem having an annular overhang and a heat shield secured to the overhang is described. More specifically, and in one embodiment, the nozzle stem includes an upstream end and a downstream end. The annular overhang is intermediate to the upstream end and the downstream end of the stem. The heat shield includes a first end and a second end, and the heat shield is welded to the annular overhang at the heat shield first end. An annular air gap is between the nozzle stem and the heat shield.

Abstract: A combustor liner includes a wall having an outboard surface and an opposite inboard surface. A plurality of first holes are inclined through the wall in a multihole pattern to channel cooling fluid therethrough to form a cooling film layer along the inboard surface. A second hole extends perpendicularly through the wall within the multihole pattern to form a shadow along the inboard surface devoid of the first holes. A transition hole extends through the wall in the shadow at a greater inclination than the first holes for cooling the wall at the shadow.

Abstract: A turbine wall includes an outer surface for facing combustion gases, and an opposite inner surface for being impingement air cooled. A plurality of adjoining ridges and grooves are disposed in the inner surface for enhancing heat transfer by the impingement cooling air.

Abstract: A nozzle failsafe nozzle actuating system for an aircraft gas turbine engine variable exhaust nozzle having a plurality of pivotal flaps circumferentially disposed about a nozzle centerline and bounding an exhaust gas flowpath in the nozzle is provided with a number of primary actuators which are operably connected to nozzle flaps for pivoting and setting an attitude of the flaps relative to the centerline and wherein each of the primary actuators is extendable between a fully extended position and a fully retracted position, and a failsafe nozzle actuating actuator for setting the attitude of each of the flaps to a failsafe attitude during a failsafe mode of the nozzle actuating system wherein the failsafe attitude is such that each of the primary actuators is at a partially retracted position between the fully extended and the fully retracted positions.

Abstract: To cool the shroud assembly in the high pressure turbine section of a gas turbine engine, high pressure cooling air is directed in metered flow to baffle plenums and thence through baffle perforations to impingement cool the rails and back surfaces of the shroud. Impingement cooling air then flows through elongated, convection cooling passages in the shroud sections and exits to flow along the shroud front surface with the main gas stream to provide film cooling. The aft rail of the shroud sections is provided with one or more cooling holes to impingement cool the annular retaining ring or C-clip retaining the shroud sections on the shroud hangers. This cooling air then travels aftward on the inboard side of the C-clip to provide convection cooling of the C-clip. In an alternative embodiment, cooling air is directed at the aft corners of the shroud base to avoid overheating.

Abstract: A turbine airfoil includes a squealer rib extending outwardly from a tip cap to define a tip cavity thereatop. A thermal insulator is disposed in the tip cavity atop the tip cap.

Abstract: A turbine engine airfoil includes a three-pass serpentine shaped cooling cavity including a leading edge chamber, an oversized intermediate chamber, and a trailing edge chamber in flow communication. The cavity further includes a combination of a plurality of ribs and a plurality of pins, a purge air swirler, a turning vane, and a metering partition which mixes the high pressure air and provides increased air cooling and uniform flow control in the airfoil. The metering partition provides a decrease of the air pressure in the trailing edge chamber allowing for an increase in the number of trailing edge slots.

Abstract: A gas turbine engine hollow airfoil includes an airfoil outer wall having width wise spaced aparted pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending longitudinally from a root to a tip. At least one internal aft flowing serpentine cooling circuit inside the airfoil has an exit that is positioned aft of the entrance so as to have a chordal flow direction afterwards from the leading edge to the trailing edge within the serpentine circuit. At least one longitudinally extending first side wall impingement chamber is in downstream fluid communication with the serpentine cooling circuit within the airfoil and positioned between one of the side walls, preferably the pressure side wall, and a first inner wall bounding the serpentine cooling circuit. Impingement cooling apertures are disposed in the first inner wall between one of the serpentine channels and the side wall impingement chamber.

Abstract: An airfoil having extended life due to reduction in stresses. The stresses are reduced by extending the thermal barrier coating below the radius between the outer band perimeter and the inner flow path surfaces and tapering the coating thickness. This additional tapered thermal barrier coating reduces the temperature gradient across a region already having high mechanical stresses resulting from geometric considerations thereby lowering thermally-induced stresses so that low cycle fatigue life is improved.

Abstract: A method and system are used to calibrate a plurality of actuators joined to an actuation ring in a control system for vectoring exhaust flaps in a vectoring nozzle. The actuators include output rods which are positioned to respective first positions at which output stroke is measured. The actuators are repositioned to different second positions, and the output stroke thereof is again measured. The stroke measurement for each of the actuators is calibrated to effect a common linear response between the first and second measurement positions. The calibrated actuators ensure coordinated linear actuation thereof as the output rods are positioned between their minimum and maximum extensions.

Abstract: An article comprising an airfoil surface is provided with a thermal insulating outer layer on at least one region, but not all, of the airfoil surface. The region is selected from the airfoil surface which has been observed to experience more strenuous environmental operating conditions during service operation than the airfoil surface outside of the region. In one form, the thermal insulating outer layer is applied to at least one region of each of a plurality of airfoil surfaces. Each airfoil surface is disposed about an axis of rotation with the region of each airfoil surface facing generally outwardly from the axis of rotation and from each other. The airfoil surfaces are rotated about the axis of rotation while a thermal insulating material contacts at least the regions.

Abstract: A gas turbine engine airfoil includes an axial serpentine cooling circuit therein. A plurality of the serpentine circuits are preferably stacked in a radial row along the airfoil trailing edge for cooling thereof.

Abstract: Methods and apparatus for suppressing cell howl with negligible impact on engine test conditions are described. The apparatus, in an exemplary embodiment includes a flow distorter configured to be positioned close to a nozzle exit of an engine nozzle, and a flow distorter support for maintaining the flow distorter at a selected location. The flow distorter is adjustably secured to the support so that a distance at which a tip of the distorter is located relative to support is adjustable.

Abstract: A turbine blade includes a hollow airfoil extending from an integral dovetail. The airfoil includes sidewalls extending between leading and trailing edges and longitudinally between a root and a tip. The sidewalls are spaced apart to define a flow channel for channeling cooling air through the airfoil. The tip is tapered longitudinally above at least one of the sidewalls and decreases in thickness.

Abstract: A turbine nozzle includes outer and inner bands between which extend a plurality of vanes for channeling combustion gases. Each of the vanes includes leading and trailing edges, and pressure and suction sides extending therebetween, and also a bow along the trailing edge to increase pressure in the gases adjacent the inner band. The vanes also include a thermal barrier coating (TBC) selectively disposed solely along the suction side between the leading and trailing edges.

Abstract: A fixed geometry duct defining a gas turbine engine exhaust nozzle includes a plurality of circumferentially adjoining flat triangular facets and coaxial first and second openings at opposite ends thereof. A method of making the duct includes scribing a plurality of lines on a flat sheet metal plate to define the facets, and bending the plate at each of the lines to form the duct.

Abstract: A turbine nozzle includes radially outer and inner bands between which extend a plurality of vanes. The vanes include impingement baffles having impingement holes for impinging cooling air inside the vane for cooling thereof. An air circuit is disposed in the inner band and includes a plurality of outlet holes. A transfer tube is disposed in flow communication between respective ones of the baffles and the air circuit for channeling thereto a portion of the cooling air from inside the baffles as pre-impingement air for discharge through the outlet holes.

Abstract: An airfoil core die includes a tab for forming a continuous and smooth contour and extension of a first trailing edge slot recessed wall to a bottom of the airfoil. The contour is formed on each airfoil core at injection and is maintained through to finish casting of the airfoil.