Abstract: Disclosed is a process of fabricating a fuel nozzle assembly, a process of fabricating a fuel nozzle ring, and a fuel nozzle ring. The process of fabricating the fuel nozzle assembly includes providing a fuel nozzle end cover, positioning a fuel nozzle ring within a cavity of the fuel nozzle end cover, and securing the fuel nozzle ring to the fuel nozzle end cover to form features corresponding to one or more of the fuel nozzle end cover and a fuel nozzle insert. The process of fabricating the fuel nozzle ring includes forming the fuel nozzle ring with at least one surface corresponding to one or more of a fuel nozzle end cover and a fuel nozzle insert. The fuel nozzle ring includes at least one surface corresponding to one or more of a fuel nozzle end cover and a fuel nozzle insert.

Abstract: A turbine is provided. The turbine includes a housing radially extending around a turbine rotor including a first piece defining a portion of a first scroll passage boundary and a second piece having an interface wall contiguous with an interface wall of the first piece, the second piece coupled to the first piece and including a divider defining another portion of the first scroll passage boundary and a portion of a second scroll passage boundary.

Abstract: A stator subassembly includes an array of circumferentially arranged stator vanes. An attachment liner secures the stator vanes to one another to provide the subassembly. A damper spring is integral with the attachment liner and is provided between the array and an outer case, which supports the array. The damper spring is configured to bias the array radially inwardly from the outer case.

Abstract: A system for sealing a gas path in a turbine includes a stator ring segment, a shroud segment adjacent to the stator ring segment, and a first load-bearing surface between the stator ring segment and the shroud segment. A first non-metallic gasket is in contact with the first load-bearing surface between the stator ring segment and the shroud segment. A method for sealing a gas path in a turbine includes placing a non-metallic gasket between any two of a stator ring segment, a shroud segment, and a casing.

Abstract: According to one aspect of the invention, a turbine assembly includes a first static structure and a second static structure radially outward of the first static structure. The assembly also includes a support member placed in a recess of the second static structure, wherein the support member includes first and second curved surfaces to contact the first and second static structures, respectively, and wherein the support member includes a biasing structure to retain the support member in the recess.

Abstract: A sealing assembly for use with a rotating machine is described herein. The rotating machine includes a stator casing that includes a radially inner surface that defines a cavity therein, and a rotor positioned within the cavity and spaced inwardly from the stator inner surface. The sealing assembly includes a support ring coupled to the stator casing, and a plurality of plate members coupled to the support ring and oriented circumferentially about the rotor. A resistance member extends inwardly from the support ring towards the rotor outer surface. The resistance member is coupled to the support ring and extends through each plate member of the plurality of plate members. At least one rotor land is defined circumferentially about the rotor outer surface. Each plate member is oriented adjacent the at least one rotor land.

Abstract: An airfoil and method of fabricating an airfoil including a first and a second side coupled together at a leading and a trailing edge and extending there between. The airfoil includes a plurality of first chord sections having a first chord length and extending outward from one of the first side or second side of the airfoil at the leading edge and a plurality of second chord sections having a second chord length and extending outward from the one of the first side or the second side of the airfoil at the leading edge. The leading edge including spaced-apart wave-shaped projections defining a waveform. The configuration defining a three-dimensional crenulated airfoil configured to facilitate desensitization of an airfoil unsteady pressure response to at least one impinging upstream generated wake or vortex by decorrelating spatially and temporally and reducing in amplitude an unsteady pressure caused by interaction of the airfoil with the upstream generated wake or vortex.

Abstract: A combustion chamber in which the injection system is attached so as to prevent any axial motion of the injection system, the combustion chamber includes: a baffle including a tubular portion having a first upstream end surrounded by a first radially projecting collar; and an injection system including a bowl which includes a cylindrical portion inserted into the tubular portion, the cylindrical portion including a second upstream end surrounded by a second radially projecting collar, the second collar axially bearing against the first collar, wherein the combustion chamber includes a clamp which axially clamps the first and the second collars against one another, and a fastener which retains the clamp clamped on either side of the first and second collars.

Abstract: A stator blade ring comprising a plurality of stator blade modules defining an annular chamber is provided. The plurality of stator blade modules comprises an elongated blade portion comprising a first and a second blade shell portion, a longitudinal passageway, and at least one opening extending through at least one of the first and the second blade shell portion to the longitudinal passageway, an inner portion brazed to a first longitudinal end of the elongated blade portion, wherein the inner portion comprises a through hole forming a portion of the annular chamber, and an inner passageway extending from the through hole to the longitudinal passageway, and an outer portion brazed to a second longitudinal end of the elongated blade portion and engaged to a steam turbine, the outer portion comprising an outer passageway open to a surface of the steam turbine and the longitudinal passageway.

Abstract: A method and structure negating thermal inertia of a solid portion, such as shielding, which influences mechanical behavior of parts of an engine structure, by reducing forces transmitted to the engine structure during an event in which a turbine blade ruptures, while preserving mechanical strength and vibratory positioning. The structure includes a flexible fastener for a shield, enabling producing a fusible section. An assembly for mounting a shielding onto a casing of the engine structure of a turbine includes connection lugs and points for attachment to the casing and shielding. The attachment points are sufficiently spaced apart in accordance with curvature of the shielding and casing so connection between the lugs and the shielding or casing, and connection of the attachment points, are substantially tangential. The lugs are sized so the connection has sufficient flexibility to control vibratory positioning thereof and ensure sufficient mechanical strength under thermo-mechanical loads.

Abstract: A stator for a turbo-machine having a plurality of airfoils extending radially therefrom has a base from which the airfoils depend, and slits disposed in the base, each slit disposed adjacent a pair of airfoils, wherein a first set of adjacent slits and a distance between a second set of adjacent slits varies

Abstract: In one aspect, a rotor hub assembly for a wind turbine is disclosed. The rotor hub assembly may generally comprise an inner hub including a plurality of hub segments assembled together and a hub enclosure including a plurality of enclosure segments assembled together. The hub enclosure may be configured to at least partially enclose the inner hub. In addition, the rotor hub assembly may include a plurality of support members extending between the inner hub and the hub enclosure.

Abstract: A turbine nozzle is constructed from a plurality of associated circumferential turbine nozzle segments each of which incorporates at least one pre-formed annular-segment passage bounded at a first azimuthal boundary by a first portion of a first bounding nozzle vane, and at a second azimuthal boundary by a second portion of a second bounding nozzle vane, with the at least one pre-formed annular-segment passage therebetween, wherein when assembled in the turbine nozzle, a first portion of a first bounding nozzle vane of one circumferential turbine nozzle segment is joined to a second portion of a second bounding nozzle vane of another adjacent circumferential turbine nozzle segment so as to form therebetween a nozzle vane of the turbine nozzle, wherein the minimum through-flow area of the associated at least one pre-formed annular-segment passage is substantially unaffected by the assembly of the circumferential turbine nozzle segments to form the turbine nozzle.

Abstract: A turbine exhaust case for a gas turbine engine has an outer shroud and an inner shroud concentrically defining therebetween an annular gaspath for channelling hot gases. A plurality of circumferentially spaced-apart turbine exhaust struts extends radially across the hot gaspath. The inner shroud may be provided in the form of an acoustic panel. The acoustic panel has a radially outwardly facing surface defining the radially inner flow boundary surface of the gaspath. The acoustic panel is mechanically fastened to the radially inner end of the struts.

Abstract: A vane segment for a gas turbine engine includes an airfoil disposed between an outer platform and an inner platform. A slot for receiving a seal is defined within the outer platform and includes closed first and second ends and an upper surface spaced apart from a lower surface with a spacing between the upper surface and the lower surfaces that varies along a length of the slot.

Abstract: A fiber preform for a turbine ring sector obtained by three-dimensional weaving and including a base-forming first portion, two tab-forming L-shaped portions each presenting two branches, two ends of the first portion being extended by respective ones of the first branches, and a second portion connecting together the two tabs, first and second strips woven together forming the first branches, a first fraction of the thicknesses of the second branches, and the first portion, there being a non-linked zone between them that is situated in the first branches and in the first portion, and a third woven strip forming the second portion and a second fraction of the thickness of the second branches of each of the tabs.

Abstract: A stator ring for use in a compressor of a gas turbine engine and a method of assembling the stator ring are provided. The stator ring includes a first plurality of segments that includes a first segment including a first circumferential end formed with a first cut and a second circumferential end formed with a second cut that is complementary to the first cut, wherein the first plurality of segments are configured to be circumferentially-coupled together, and a second plurality of segments that includes a second segment including a first circumferential end with a third cut and a second circumferential end with a fourth cut that is complementary to the third cut, wherein the second plurality of segments are configured to be circumferentially-coupled together.

Abstract: A ram air fan inlet housing for containing a ram air fan rotor. The inlet housing includes a flanged surface and an interior surface. The flanged surface is perpendicular to an axis of the inlet housing and defines a flange plane at an axial end of the inlet housing. The interior surface is symmetric about the axis of the inlet housing and includes a flange section, a transition section, an outlet section, a rotor section, and an inlet section.

Abstract: A ram air fan inner housing for a ram air fan assembly comprises a center body housing, an end cup attached to the center body housing, and a perforated cone. The perforated cone is attached to the center body housing and the end cup such that the perforated cone extends away from the center body housing and radially inward toward an axis of the inner housing.

Abstract: A diffuser for a ram air fan assembly includes a perforated cone, an inlet ring seal, and an outlet ring seal. The perforated cone has a frustoconical shape symmetrical about an axis of the diffuser. The inlet ring seal is attached to, and axially disposed about, a first end of the perforated cone. The inlet ring seal includes a fan housing connection having a cylindrical shape. The outlet ring seal is attached to, and axially disposed about, a second end of the perforated cone. An average external diameter of the second end is greater than an average external diameter of the first end such that the perforated cone extends away from the inlet ring seal and radially outward from the axis of the diffuser.

Abstract: The present invention relates to a method for manufacturing a turbomachine rectifier comprising an outer ring (5) and a plurality of stator vanes (2), said vanes comprising a platform (9) and a blade (12), wherein said method comprises at least the following steps: a) first folds (7) are draped over a core comprising perforations and having the shape of the aerodynamic flow; b) optionally, the first folds (7) are cut along incisions (8) placed opposite the perforations of the core; c) the blades (12) of the vanes or, alternatively, the blades (12) of the vane preforms, are inserted through the incisions (8) of the first folds (7) and through the perforations of the core; d) last folds (10) are draped over the platforms (9) thus forming a rectifier preform; e) a resin is injected into a closed mold with the preform and the resin-impregnated preform is polymerized; f) a molded piece, essentially having the shape and dimensions of said rectifier, is retrieved from the mold.

Abstract: The present application thus describes a method of manufacture for a late lean injection system in a combustor of a combustion turbine engine.

Abstract: An arrangement is disclosed in which deflections and mechanical stresses in turbine low pressure casings are controlled by inserting struts between the turbine diaphragm ledge ring stages. Preferably, the struts are cylindrical in shape, although other different shapes can be used in accordance with the needs of different applications. The struts can be solid or hollow in construction, although preferably they are hollow to reduce the material needed to fabricate them. The struts are located around the turbine casing, as is required to connect together two ledge rings where there is high axial deflection in the casing.

Abstract: A seal is provided that includes a resilient, arc-shaped body having a pair of sealing portions movable between a pre-compressed state and a compressed state. A removable pre-compression member slidingly engages the seal body along at least a portion of its length to hold the sealing portions in the pre-compressed state. In an embodiment, the seal body includes a pair of first retaining features and the pre-compression member includes a retaining portion having a pair of second retaining features that engage the pair of first retaining features to hold the sealing portions in the pre-compressed state. A strap portion extends from and between each retaining portion to allow the pre-compression member to flex during assembly of the seal to accommodate the radius of the seal body.

Abstract: A combustor (100) for a gas turbine engine (30) has a circumferentially extending liner (109) defining at least a portion of an interior combustion chamber (107) and a hot gas path (115). The liner includes a resonator section (112) including at least one resonator (116A, 116B) having a resonator chamber (125A, 125B) formed on an exterior of the liner. A thermal barrier coating (118) is disposed along an inner surface of the liner including an inner surface (130) of the resonator section. The resonator further includes a plurality of apertures (117) and each aperture extends through the liner and the thermal barrier coating at the resonator section and there is fluid flow communication between the combustion chamber and the resonator chamber.

Abstract: A turbine nozzle for an air cycle machine includes a base with a multiple turbine vanes which each extend for a vane height H, a throat width W defined between each of the multiple of turbine vanes, wherein a ratio W/H is 0.101-0.112.

Abstract: A piping assembly and a method for connecting an inner shell and an outer shell in a turbine system are disclosed. The piping assembly includes an inner fitting for connection to the inner shell and an outer fitting for connection to the outer shell. The piping assembly further includes an attenuation member extending between the inner fitting and the outer fitting and comprising an attenuation curve. The attenuation curve permits movement of the attenuation member in at least one of a longitudinal direction, a radial direction, or a tangential direction.

Abstract: A method of fabricating a component is provided. The component includes a substrate that has at least one interior space. The method includes forming one or more grooves in the component. Each groove extends at least partially along an outer surface of the substrate and narrows at an opening thereof, such that each groove is re-entrant shaped. A cross-sectional area A of each groove is in a range of about 2 to about 3 times an area R=W*D, where W is the width of the opening and D is the depth of the re-entrant-shaped groove. Components with grooves formed in the substrate and components with grooves formed at least partially in a structural coating are also provided.

Abstract: In certain embodiments of the present disclosure, a turbine casing assembly is described. The turbine casing assembly includes an inner casing and an outer casing surrounding the inner casing. The outer casing has a first outer casing section and a second outer casing section that join together along a flange. A bolt extends through the flange and joins together the first outer casing section and the second outer casing section. The turbine casing assembly further includes a pin. The pin has a segment defining an opening therethrough. The pin extends through the inner casing and the outer casing and supports the inner casing relative to the outer casing. The bolt passes through the opening defined by the pin.

Abstract: A method of manufacturing a multistage exhaust turbocharger having a high-pressure stage turbocharger and a low-pressure stage turbocharger includes providing a high-pressure turbine housing of the high-pressure stage turbocharger, the high-pressure turbine housing being integrally formed with the exhaust manifold by casting or welding, assembling constituent parts of the high-pressure stage turbocharger using the high-pressure turbine housing as a reference, after the assembling constituent parts of the high-pressure stage turbocharger, attaching the low-pressure stage turbocharger to the high-pressure stage turbocharger by attaching a low-pressure turbine housing to a low-pressure turbine connection flange, and after the attaching the low-pressure stage turbocharger, attaching an air supply channel to connect between a supply air outlet of the low-pressure stage turbocharger and a supply air inlet of the high-pressure stage turbocharger, wherein the low-pressure turbocharger is disposed below the high-press

Abstract: A supersonic inlet includes a relaxed isentropic compression surface to improve net propulsive force by shaping the compression surface of the inlet to defocus the resulting shocklets away from the cowl lip. Relaxed isentropic compression shaping of the inlet compression surface functions to reduce the cowl lip surface angle, thereby improving inlet drag characteristics and interference drag characteristics. Supersonic inlets in accordance with the invention also demonstrate reductions in peak sonic boom overpressure while maintaining overall engine performance.

Abstract: A tie rod includes a gusset which extends between a rod and a base.

Abstract: An example turbomachine load management assembly includes a support rod extending from a first rod end to a second rod end that is moveable radially relative to a fan duct, a bearing housing, or both, when the support rod is coupling the fan duct and the bearing housing.

Abstract: A fuel nozzle is provided and includes a nozzle body defining first and second interior regions for providing a supply of first and second fluids, a collar defining a third interior region and radial slots permitting radial ingress of a third fluid to the third interior region and a nozzle tip interposed between the nozzle body and the collar. The nozzle tip defines an annular slot, first discrete passageways by which the first fluid is communicated from the first interior region to the annular slot, second discrete passageways by which the first fluid is communicated from the annular slot to the radial slots, and third discrete passageways by which the second fluid is communicated from the second interior region to the radial slots.

Abstract: A blade of a turbomachine, in particular of a gas turbine, is disclosed. The blade includes at least one fastening section for fastening the same in a locating section of a rotor-side or stator-side component of the turbomachine. The, or each, fastening section has a core region which is made of a relatively brittle and relatively light material and is embedded in an encasing region made of a material having relatively high ductility.

Abstract: A fixture for assembling a stator vane assembly of a gas turbine engine includes a base which defines a multiple of locators to position each of a multiple of vanes relative to an outer fairing and an inner fairing. A first tangential wedge clamp is receivable in a first aperture located within the base for each of the multiple of vanes. A second tangential wedge clamp is receivable in a second aperture located within the base for each of the multiple of vanes.

Abstract: A system for supporting a shroud used in an engine has a shroud positioned radially outboard of a rotor, which shroud has a plurality of circumferentially spaced slots; a forward support ring for supporting the shroud; the forward support ring having a plurality of spaced apart first tabs on a first side for functioning as anti-rotation devices; the forward support ring having a plurality of spaced apart second tabs on a second side; and the second tabs engaging the slots in the shroud and circumferentially supporting the shroud.

Abstract: A gas turbine engine mid turbine frame apparatus includes a an inner case supporting a bearing, an outer case to which the inner case is mounted, and an apparatus for radially positioning one relative to the other, the apparatus including a plurality of radial locators extending outwardly of the outer case and an associated locking apparatus. The locking apparatus is mounted adjacent to but independent of the locators and provides for anti-rotation of the radial locators once installed.

Abstract: A vane assembly (10) having: an airfoil (12) and a shroud (14) held together without metallurgical bonding there between; a channel (22) disposed circumferentially about the airfoil (12), between the airfoil (12) and the shroud (14); and a seal (20) disposed in the channel (22), wherein during operation of a turbine engine having the vane assembly (10) the seal (20) has a sufficient ductility such that a force generated on the seal (20) resulting from relative movement of the airfoil (12) and the shroud (14) is sufficient to plastically deform the seal (20).

Abstract: A method facilitates extending the useful life of a turbine engine including a combustor and a turbine downstream from and in flow communication with the combustor. The method comprises identifying at least a portion of the combustor that is at least one of distressed, worn, deteriorated, satisfied pre-determined threshold data, and beyond serviceable limits, identifying at least one enhancement kit for installation within the combustor, and removing at least two portions of the combustor including at least two of a combustor dome assembly, a swirler/cowl assembly, and the combustor inner and outer liners. The method also comprises installing the enhancement kit within the turbine engine such that a useful life of the turbine engine is facilitated to be increased.

Abstract: A combustor dome heat shield and a louver are separately metal injection molded and then fused together to form a one-piece combustor heat shield.

Abstract: A method of forming and/or assembling a multi-panel outer wall (14) for a component (12) in a machine subjected to high thermal stresses comprising providing such a component (12) that includes an inner panel wall (16) having an outer surface, and an array of interconnecting ribs (38) on the outer surface of the component (12). An intermediate panel (22) is provided and preferably preformed to a general outer contour of the component (12), and is positioned over the inner panel (16). An external pressure force is applied across a surface area of the intermediate panel (22) against the outer surface of the component (12) to contour the intermediate panel (22) according to a geometric configuration formed by the ribs (38) thereby forming cooling chambers (24) between the outer surface and ribs (38) of the component (12) and the intermediate panel (22).

Abstract: A head part (100) of an annular combustion chamber for a gas turbine, the head part comprising: an end wall (110) with a first opening (112) for accommodating a burner; a heat shield (120) covering a back side of the end wall which faces towards the combustion chamber, the heat shield comprising a second opening (124) for accommodating the burner; and a burner collar (130) adapted to fit within the second opening and receive the burner; wherein the head part of the annular combustion chamber is configured such that in an installed configuration the burner collar is fastened to the heat shield with one or more first fasteners (142) to form a sub assembly (140) and the heat shield of the sub assembly is fastened to the end wall with one or more second fasteners (144).

Abstract: A head part of an annular combustion chamber for a gas turbine, the head part comprising: an end wall with a passage opening for accommodating a burner, the end wall comprising a back side which faces towards the combustion chamber; a burner collar adapted to fit within the passage opening and receive the burner, the burner collar comprising a protruding portion radially protruding from an outer surface of the burner collar; and a ring member for receiving the protruding portion of the burner collar; wherein the head part of the annular combustion chamber is configured such that in an installed configuration a radially outer portion of the ring member engages with the end wall and a radially inner portion of the ring member receives the protruding portion of the burner collar.

Abstract: A rotor (1) of a gas turbine compressor comprises a multiplicity of welded-together rotor disks (3a, 3b, 4, 5), of which two or more rotor disks (3a, 3b), in a radially outer region (9?), are welded together, and, in a radially inner central region (9), are abutment-joined together. Via the abutment-joining of two rotor disks (3a, 3b), a heat flow (8) radially outward from the center of the rotor (1) is achieved so that the material temperature of the rotor (1) can be kept below a predetermined level during operation. As a result, the service life of the rotor (1) can be increased. In one embodiment, one rotor disk (3a), on its surface, additionally has a recess (7) which can be cooled from outside. The rotor disks (3a, 3b) according to the invention, which are welded and abutment-joined together, can especially be used at the last point in the flow direction of the compressor.

Abstract: A method for repairing a turbine nozzle segment includes providing a turbine nozzle segment having an enclosure with an integral discourager, removing the enclosure and/or the discourager from the turbine nozzle segment and attaching a new enclosure ad/or discourager to the turbine nozzle segment.

Abstract: One embodiment of the present invention is a unique gas turbine engine case. Another embodiment is a unique method of manufacturing a gas turbine engine case. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engine cases. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A stem (34) extends from a second part (30) through a hole (28) in a first part (22). A groove (38) around the stem provides a non-threaded contact surface (42) for a ring element (44) around the stem. The ring element exerts an inward force against the non-threaded contact surface at an angle that creates axial tension (T) in the stem, pulling the second part against the first part. The ring element is formed of a material that shrinks relative to the stem by sintering. The ring element may include a split collet (44C) that fits partly into the groove, and a compression ring (44E) around the collet. The non-threaded contact surface and a mating distal surface (48) of the ring element may have conic geometries (64). After shrinkage, the ring element is locked onto the stem.

Abstract: The stationary vane unit of a rotary machine includes: a first band member that extends in the circumferential direction and comes into contact with the outer shrouds of the plurality of stationary vane members from one side thereof in the main axial direction in which a central axis extends; a second band member that extends in the circumferential direction and comes into contact with the outer shrouds of the plurality of stationary vane members from the other side thereof in the main axial direction; and a fastening member that fastens the first band member and the second band member to each other so that the outer shrouds of the plurality of stationary vane members are connected to each other.

Abstract: An apparatus is provided for coupling a first portion of a gas turbine transition duct to a second portion of a gas turbine transition duct to reduce vibratory deflection. The apparatus may comprise: at least one first support structure attached to the gas turbine transition duct first portion; at least one second support structure attached to the gas turbine transition duct second portion; and at least one coupling mechanism configured to couple the at least one first support structure to the at least one second support structure so as to allow sliding movement between the at least one first support structure and the at least one second support structure when a movement force of the at least one first support structure and the at least one second support structure exceeds a predefined frictional force threshold value.