Abstract: A turbine nozzle assembly includes a plurality of nozzle segments and a nozzle support for supporting the nozzle segments. Each nozzle segment includes an outer band, an inner band and at least two vanes disposed between the outer and inner bands. A retention flange extends radially inwardly from the inner band and has a first hole formed therein. The nozzle support includes a recess formed therein and a mounting flange extending therefrom. The mounting flange is disposed in contact with the retention flange and has a second hole formed therein. A pin is disposed in the first and second holes to position the flanges with respect to one another. A pin retainer is disposed in the recess and has a holding flange for retaining the pin in place. The nozzle support includes a substantially conical portion and an air seal integrally formed thereto.

Abstract: A shroud panel for a turbine shroud includes forward and aft hooks which are used to support the panel radially above a plurality of turbine rotor blades. The panel forward hook has radially outer and inner lands, with the outer land being defined by a plurality of pads circumferentially spaced apart from each other by a respective recess. The panel forward hook is sized to engage the complementary forward slot of the turbine shroud substantially concentrically therein. The pads and recess restrict flow leakage around the panel forward hook for maintaining backflow margin and improving clearance control.

Abstract: A turbine shroud includes an annular outer casing having first and second radial flanges between which are suspended an adjoining annular shroud support and an annular shroud hanger. A plurality of shroud panels are joined to the shroud hanger and are positionable radially above turbine rotor blades to define a tip clearance therebetween. The four flanges are disposed in abutting contact with each other, and respective pluralities of circumferentially spaced apart first, second, and third recesses are disposed respectively between the casing first flange and the shroud flange, between the shroud flange and the hanger flange, and between the hanger flange and the casing second flange for providing radial flowpaths therebetween.

Abstract: A turbine shroud hanger is supported from an annular outer casing and includes an annular radial flange and integral forward and aft legs at a radially inner end thereof. The legs extend axially oppositely to each other and include respective distal ends configured for supporting a plurality of arcuate shroud panels radially above a plurality of turbine rotor blades to define a tip clearance therebetween. The legs include circumferentially spaced apart forward and aft slots extending axially from the distal ends thereof toward the radial flange, and completely radially therethrough to bifurcate the legs into circumferential segments for reducing transient thermal expansion of the hanger to reduce in turn expansion of the tip clearance.

Abstract: A turbine shroud includes an annular shroud support joined to an annular outer casing and spaced radially inwardly therefrom to define an annular flow duct for receiving compressor bleed air. A plurality of shroud panels are joined to the shroud support and have a radially inner surface positionable radially above turbine rotor blades to define a tip clearance therebetween. An isolation shield is joined to the shroud support radially above the panels and includes a plurality of open-ended cells facing radially outwardly in the duct toward the outer casing for inducing drag in the bleed air flowable axially through the duct to thermally isolate the shroud support from the bleed air for controlling the tip clearance.

Abstract: A turbine cooling supply circuit for a gas turbine engine is disclosed in which the flow of coolant through the engine is directed to minimize temperature rise prior to discharge into the turbine. Bleed air from an impeller stage of a compressor is channeled along a backwall thereof into a cavity disposed radially inwardly of a combustor casing. The cavity is divided by a lightweight tubular member having a circumferentially uniform contour into a non-flow zone proximate the casing and a flow zone proximate a rotor. Maintenance of a high tangential flow component in the bleed flow in the flow zone facilitates discharge of the flow through apertures in a rotor into a bore portion thereof, minimizing work input and temperature rise of the coolant.

Abstract: A turbine shroud hanger includes forward and aft radially outer hooks for mounting the hanger to an annular shroud support, and forward and aft radially inner hooks for supporting a shroud panel radially above a plurality of turbine rotor blades for controlling tip clearance therebetween. The hanger forward outer hook has a radially outer land defined by a pair of pads at distal ends of the hook, with a ridge extending circumferentially therebetween. The outer land is interrupted by a central recess extending between the pads and aft from the ridge to a distal edge of the hook. The forward outer hook has a predetermined radial height which is predeterminedly larger than the radial height of the corresponding forward slot in the shroud support for effecting an interference fit at least in part at the pair of pads to provide a seal against leakage of bleed air.

Abstract: A turbine disk cooling system for a gas turbine engine of a type having a compressor section including first and second disks connected by a seal and defining a compressor interstage volume, a turbine section including first and second stage disks connected by a seal and defining a turbine inner stage volume, and a compressor shaft interconnecting the compressor disks with the turbine disks including orifices located in the compressor seal for allowing air in the compressor section to enter the compressor interstage volume, and a radial inflow impeller for guiding the compressor air towards the compressor shaft. The impeller including a plurality of radially-extending conduits suspended in the compressor interstage volume, the conduits having a plurality of interior, radially-extending passages for conveying the compressor air from the orifice to the compressor shaft.

Abstract: A coated article comprises a base article having a substrate made of a material selected from the group consisting of a nickel-base alloy and a cobalt-base alloy, an intermediate metallic coating structure, and a thermal barrier coating. The intermediate metallic coating structure may be a protective layer of platinum, rhodium, or palladium on the base article, with no bond coat. The intermediate metallic coating may also be a bond coat, either oxidized or unoxidized, with a protective layer of platinum, rhodium, or palladium thereon. The protective layer typically partially interdiffuses with the underlying metallic structure.

Abstract: Cooling of a workpiece is conducted at selected workpiece locations in accordance with predetermined heat flux, at the locations, required to result in a desired workpiece cooling rate for workpiece integrity, microstructure and mechanical properties. A cooling fluid is controlled to follow the workpiece surface according to preselected cooling fluid convective cooling parameters including, but not limited to, cooling fluid direction, mass flow rate, and velocity at the selected locations.

Abstract: There is provided by the present invention method and apparatus for damping vibrations or other unwanted motions of a brush seal, the apparatus including a damper having a plurality of individual plates that are bonded directly to the inlet rows of bristles of the seal. Methods for producing a damped brush seal in accord with the present invention contemplate brazing damping plates to the bristles, absorbing a melting point suppressor into the surface of the damping plates and then heating the damping plates until the treated surface melts and bonds with the bristles, or sintering a braze alloy to the bristles such that the braze alloy functions as a damping plate.

Abstract: A rotor disk assembly includes rotor disks and structural spacer arms adapted to transmit axial loads and bending moments between adjacent disks. The spacer arms include cooling fins for convective cooling of the spacer arms. The spacer arm is preferably a self supporting wheel structure. The cooling fins are preferably circumferentially continuous and extend radially into a relatively cooler rotor bore cavity in which the disk hubs are supported. The cooling fins can be positioned on the spacer arms to reduce thermal distortion of the spacer arms, and thereby reduce bending stresses transmitted to the disks. The spacer arm includes a body section adjacent a relatively higher temperature seal cavity, which can be purged by cooling air from the relatively lower temperature rotor bore cavity.

Abstract: A circumferential element discourages the flow of hot gases exiting from a labyrinth or other type of seal in a gas turbine engine. This seal exit flow discourager may be integrally incorporated in the honeycombed seal surface of a labyrinth seal or in the ring supporting the honeycombed seal surface. The seal exit flow discourager has a generally radial surface upon which impinges the jet of hot gases exiting from the labyrinth seal. This surface prevents the exiting hot gases from impinging directly on the low-pressure turbine shaft. At the same time low-temperature air from the compressor is flowing through an annular passageway located between the labyrinth seal and the low-pressure turbine shaft. The seal exit flow discourager diverts the hot gases exiting from the labyrinth seal into the stream of cooling air, thereby causing turbulent mixing of the hot gases and cool air.

Abstract: An interference fit up for an HPT hanger is accomplished by machining and casting features into the hanger which create a spring effect when the hanger is assembled into the support. The spring effect may be accomplished in a variety of ways, including offset radial cut features on the hanger relative to the support, or offset projection features on the hanger which are concentric with the support. The spring effect is then accomplished by a flex or deflection of the ends of the hanger to conform to the width of the support, and clearance indentations created during the installation.

Abstract: A thermal control apparatus is provided for gas turbine engine cases and particularly for thermal control of clearances between turbine rotors and surrounding shrouds. Thermal control of an annular case section or ring is provided by counterflowing two heat transfer fluid flowpaths in heat transfer communication with the section of engine casing. The flowpaths may be in parallel or series such that there is substantially no circumferential gradient in the mass flowrate weighted average temperature of the heat transfer fluid supplied by the two counterflowing fluid flowpaths. One embodiment provides forward and aft rings cooled by three spray tubes in each of two 180.degree. sectors whereby the forward and aft spray tubes flow thermal control air flowing in one circumferential direction and the middle spray tubes flows thermal control air in an opposite circumferential direction.

Abstract: A heat shield mechanism for thermally protecting a casing located in a turbine engine having a plurality of honeycomb cells which are connected to a support plate. A spring in contact with the support plate and in contact with a vane liner exerts a force on the support plate which causes at least one of the plurality of honeycomb cells to be pressed against the casing.

Abstract: A heat shield mechanism for shielding a structure in a turbine engine having a sheet metal backing which is bonded to a plurality of honeycomb cells aligned in a radially outward manner toward a casing. The honeycomb cells and sheet metal backing form a heat shield which acts as a vortex destroyer by eliminating fluid flow around the turbine structure.

Abstract: To cool the shroud 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 shroud rails and back surface. Impingement cooling air then flows through elongated, convection cooling passages in the shroud and exits to flow along the shroud front surface with the main gas stream to provide film cooling. The baffle perforations and the convection cooling passages are interactively located to achieve maximum cooling benefit and highly efficient cooling air utilization.

Abstract: A heat shield device for a vane liner in which a first plurality of honeycomb cells which are axially aligned are attached to a first side of the vane liner. A second plurality of honeycomb cells are axially aligned and attached to a second side of the vane liner. The first and second sides of the vane liner are opposite to each other and are provided with cavities which are filled in by the first and second pluralities of honeycomb cells.

Abstract: Apparatus for statically determinant mounting of a vane assembly in a gas turbine engine includes means for assuring contact support of the vane assembly in a predetermined number of identified locations. The vane assembly includes at least one vane extending between a radially outer nozzle band and a radially inner nozzle band. The radially outer nozzle band has axially directed load bearing elements on distal ends. The radially inner nozzle band has a radially inwardly oriented flange depending therefrom. The engine includes a radially outer support member positioned to engage the load bearing elements in the radially outer nozzle band when the vane assembly is in an operative position. A radially inner support member is incorporated in the engine for engaging the flange to axially restrain the vane assembly. The inner support member includes means for releasably connecting the flange thereto for supporting the vane assembly in radially and circumferential directions.