Abstract: A diffuser for use in a gas turbine engine, the diffuser having a first wall, a second wall, and a flow separator. The first and second wall define an annular cavity, with the annular cavity having an inlet. The first and second wall also forming a prediffuser that is proximate the inlet, and a dump region distal the inlet. The flow separator extends from the first wall into the annular cavity.

Abstract: An annular combustion chamber for a turbomachine presenting an axial direction, a radial direction, and an azimuth direction, the combustion chamber including a first annular wall and a second annular wall, each annular wall defining at least a portion of an enclosure of the combustion chamber. The first annular wall and the second annular wall present complementary assembly mechanisms that co-operate by engagement in azimuth.

Abstract: A nozzle blade and nozzle ring assembly includes a nozzle blade having radially inner and outer sidewalls with an airfoil portion extending therebetween; the inner and outer sidewalls walls formed with axially-extending first surface features along forward and aft marginal edges of the inner and outer sidewalls, respectively; and radially inner and outer nozzle rings formed with corresponding axially-extending second surface features mated with the first surface features, wherein the radially inner and outer sidewalls are welded to the radially inner and outer nozzle rings only along the mated first and second surface features.

Abstract: An assembly includes a plurality of vanes, a forward liner segment, and an aft liner segment. The forward liner segment and the aft liner segment are mounted to the plurality of vanes and each segment comprises an arc of less than 360° in length.

Abstract: A wrenching mechanism includes a multiple of rectilinear fastener plates each with a nut-shaped opening, each of the multiple of rectilinear fastener plates located within one of the multiple of rectilinear openings. A fastener mounted at least partially through the nut-shaped opening of each of the multiple of rectilinear fastener plates.

Abstract: A turbine nozzle attachment assembly includes an outer turbine component (a shroud or a turbine shell) formed with a circumferential groove open in a forward-facing axial direction; a nozzle segment including a vane extending between inner and outer bands, the outer band provided with an upstanding annular hook formed with a hook element extending in an aft-facing axial direction and received in the circumferential groove. The upstanding annular hook and hook element are formed with a circumferentially-oriented slot. An anti-rotation block is located in the circumferentially-oriented slot, and an anti-tipping plate having a circumferential width greater than a corresponding circumferential width of the circumferentially-oriented slot substantially covers a forward face of the anti-rotation block. The anti-rotation block and the anti-tipping plate are fastened directly to the outer turbine component.

Abstract: A transition duct for use in a turbine engine is provided. The transition duct includes a radially inner wall and a radially outer wall positioned about the radially inner wall defining a flow passage therebetween. The radially outer wall extends and is contoured from an upstream end to a downstream end of the transition duct. As such, the slope of the radially outer wall increases from the upstream end to a predetermined axial location and decreases from the predetermined axial location to the downstream end.

Abstract: An example method of assembling a turbomachine airfoil array includes, among other things, securing a partial airfoil array within a fixture, the partial airfoil array having at least one existing airfoil extending radially between an inner and an outer fairing and an open area where at least one existing airfoil has been removed. The method includes mounting a positioning saddle relative to a base of the fixture, the positioning saddle aligned with the open area, holding a replacement airfoil using the positioning saddle, applying a curable material at an interface between the replacement airfoil and the inner and outer fairing, and curing the curable material while maintaining a relative position between the replacement airfoil and the inner and outer fairing.

Abstract: A disclosed structural guide vane for a gas turbine engine includes a strut defining a plurality of channels and a cover attached over the plurality of channels. The strut includes a first side defining a first surface of an airfoil and a second side including a plurality of channels. At least one of the plurality of channels defines a wall thickness of a portion of the first side that varies in a chord wise direction.

Abstract: A mixer of a bypass turbine aeroengine according to one embodiment, includes circumferential inner and outer flow surfaces in a wavy configuration to form a plurality of lobes of the mixer. The mixer has an upstream end portion of sheet metal with a first thickness and a downstream end portion of sheet metal with a second thickness less than the first thickness.

Abstract: In accordance with an exemplary embodiment, a method of manufacturing a stator airfoil assembly includes forming an interior wall of a stator airfoil using an additive manufacturing technique, forming an exterior wall of the stator airfoil using the additive manufacturing technique, and forming a plurality of internal ribs between the interior wall and the exterior wall using the additive manufacturing technique. A cooling air circuit is formed in a space between the interior wall and the exterior wall. Further, the interior wall, the exterior wall, and the internal ribs are formed simultaneously as an integral structure by using the additive manufacturing technique.

Abstract: A stator assembly for a rotary machine having a rotor arranged to rotate about an axis in use. The stator assembly has a circumferential support member or casing arranged about said axis and a plurality of elements extending in a substantially radial direction from the support. The elements have a platform at an end thereof for engagement within the support, wherein the elements each comprise a hollow internal cavity having an opening through the platform at the end of the element, wherein said internal cavity is filled with a vibration damping material. The elements may be filled vanes in a gas turbine engine compressor. The platforms may also be filled with the vibration damping material.

Abstract: The present invention relates to a method for replacing an inner disk element of a blade integrated disk wherein the blade integrated disk has a plurality of recesses in the form of axial through-bores in the joining area. The method begins with filling the axial through-bores, after which the inner disk element of the blade integrated disk is separated from the remaining blade ring. Next a new inner disk element is inserted into the remaining blade ring and the new inner disk element and the remaining blade ring are joined. Finally the axial through-bores are restored.

Abstract: In a method of making a turbocharger housing surfaces of a sheet-metal shell which bound a gap towards the turbine wheel are measured, after the sheet-metal shell has been joined with a bearing flange provided for support of a turbine wheel, to generate measuring values. Subsequently, the bearing flange is machined in dependence on the measuring values for establishing a position of the turbine wheel in relation to the bearing flange and thus of the sheet-metal shell.

Abstract: A gas turbine device includes an inlet casing and a stator diaphragm provided inside the inlet casing. The stator diaphragm has an integral inner stator shroud and an outer stator shroud and a plurality of stator vanes provided in a circumferential arrangement between the inner stator shroud and the outer stator shroud. A plurality of key slots is provided in a circumferential arrangement on the stator diaphragm and a corresponding plurality of key slots is provided in a circumferential arrangement on the inlet casing. The key slots provided on the stator diaphragm are aligned with the plurality of key slots provided on the inlet casing. A key is inserted into each of the plurality of key slots provided to prevent rotation of the stator diaphragm with respect to the inlet casing. The stator diaphragm is secured in an axial direction by a stator shear ring.

Abstract: A turbine engine includes multiple tierod supports. Each of the tierod supports is connected to an inner frame case of the turbine engine via multiple fasteners and at least one self anti-rotating bushing component.

Abstract: A hanger for suspending a liner within an exhaust duct of a gas turbine engine comprises a bracket, a washer, a rod and a cap. The bracket is for connection to an exhaust duct liner of a gas turbine exhaust system. The washer is connected to the bracket at a slip joint. The rod is connected to the washer at a ball joint. The cap for is for connecting to an exhaust duct of the gas turbine exhaust system. The cap is connected to the rod.

Abstract: A fuel-air mixer for use with a combustor is provided. The fuel-air mixer includes an outer ring, an intermediate hub disposed coaxially within the outer ring such that a first plenum is formed therebetween, and an inner hub disposed coaxially within the intermediate hub such that a second plenum is formed therebetween. A first swirler including a plurality of first swirler vanes is positioned between the outer ring and the intermediate hub, wherein the first swirler is configured to rotate fluid flowing therethrough in a first direction. A second swirler including a plurality of second swirler vanes is positioned between the intermediate hub and the inner hub, wherein the second swirler is configured to rotate fluid flowing therethrough in a second direction. A plurality of fuel injection conduits are also defined within the fuel-air mixer for channeling fuel to the first and second plenums.

Abstract: An exemplary turbine engine assembly includes a first shaft that is rotatably driven by a second shaft of a gas turbine engine and a compressor hub driven by the first shaft. The compressor hub is within a compressor section of the gas turbine engine. An epicyclic gear train is driven by the first shaft. A common attachment point secures the first shaft and the compressor hub to the second shaft.

Abstract: A combustor dome and heat-shield assembly comprises a heat-shield having a first opening therethrough. A swirler extends through the first opening and captures the heat-shield. A dome having a second opening therethrough and has an upstream side and a downstream side. The swirler extends through the second opening from the downstream side to the upstream side to capture the heat-shield on the downstream side, and a retaining clip engages the swirler to secure the swirler and the heat-shield on the dome.

Abstract: A fuel nozzle assembly for use with a combustor is provided. The fuel nozzle assembly includes a fuel nozzle including a discharge end. A cap is coupled adjacent to the nozzle discharge end, wherein the cap includes an outer surface. At least one dampener mechanism is coupled to the cap outer surface to facilitate reducing vibrations induced to the fuel nozzle.

Abstract: A combustion system including a combustor; a combustor liner disposed within the combustor is provided. At least one primary fuel nozzle is provided to provide fuel to a primary combustion zone disposed proximate to the upstream end of the combustor liner. A transition duct is coupled to the downstream end of the combustor liner. A secondary nozzle assembly is disposed proximate to the downstream end of the combustor to provide fuel to a secondary combustion zone at locations predetermined to reduce peak thermal loads on the surface area of the transition duct.

Abstract: A turbomachine combustor nozzle includes a first plate member having a first plurality of openings, and a second plate member having a second plurality of openings that are configured and disposed to be in alignment with the first plurality of openings. A plurality of nozzle members extends through corresponding ones of the first plurality of openings and the second plurality of openings. Each of the plurality of nozzle members includes a solid core.

Abstract: A turbomachine combustor nozzle includes a monolithic nozzle component having a plate element and a plurality of nozzle elements. Each of the plurality of nozzle elements includes a first end extending from the plate element to a second end. The plate element and plurality of nozzle elements are formed as a unitary component. A plate member is joined with the nozzle component. The plate member includes an outer edge that defines first and second surfaces and a plurality of openings extending between the first and second surfaces. The plurality of openings are configured and disposed to register with and receive the second end of corresponding ones of the plurality of nozzle elements.

Abstract: An assembly for use in a gas turbine engine includes an inner diameter (ID) shroud having a plurality of ID slots and an outer diameter (OD) shroud having a plurality of OD slots. The OD shroud is positioned radially outward from the ID shroud. A first stator airfoil extends from the ID shroud to the OD shroud. The first stator airfoil is positioned at least partially in one of the ID slots and is positioned at least partially in one of the OD slots. A first end of the first stator airfoil has one of a raised tab and an indented notch.

Abstract: A ring segment for a blade outer air seal which includes an array of separated floating wall surfaces each extending from a pedestal and having an inner side with a fibrous metal pad secured and a TBC secured over the pads. The pedestals are formed by diffusion chambers formed between an outer side and an inner side of the ring segment. Purge air holes connect the diffusion chambers to the impingement cavity formed above the ring segment. Thin slots form the separated floating wall surfaces and connect the diffusion chambers to discharge cooling air through the sides of the floating wall surfaces. Cooling air holes also pass through the pedestals to discharge cooling air onto the TBC surfaces.

Abstract: A double-walled turbocharger housing made of sheet metal includes an inner sheet metal shell which conducts hot exhaust gases and a relatively cold outer sheet metal shell which is spaced apart from the inner sheet metal shell via an air gap. The double-walled turbocharger housing is connected to a bearing flange with a circumferential collar via a common welding seam. The bearing flange is provided with a collar on an outside which faces away from the hot exhaust gas. An outer end of the inner sheet metal shell is provided with a collar with which it embraces the outer sheet metal shell. Both sheet metal shells are connected with the collar of the bearing flange via a common welding seam, wherein an air gap is present between the inner sheet metal shell and the outer sheet metal shell up to the outer end of the inner sheet metal shell.

Abstract: A turbine assembly can include a turbine wheel, a shroud component, a turbine housing and a seal that includes a wall and a lower lip that that extends radially outwardly from the wall at an obtuse angle where the seal is disposed, at least in part, between an outer surface of the shroud component and an inner surface of the turbine housing. Various other examples of devices, assemblies, systems, methods, etc., are also disclosed.

Abstract: A turbine engine stator stage includes a plurality of vanes with each of the plurality of vanes having a camber angle. The plurality of vanes is arranged in a plurality of groups with each group including a pre-determined sequence of vanes. The ordering of vanes within each group is determined by the camber of the individual vanes. This results in an arrangement of vanes within the stator stage which can modify the flow characteristics of the air entering the stator stage to reduce the circumferential pressure variation in the flow region immediately downstream of the stator stage.

Abstract: Disclosed herein are apparatuses, methods, and systems for cross sectional area convergence reduction and selection. In an embodiment, an aft end cross sectional area is determined for a transition piece and a forward end cross sectional area is determined for the transition piece. The transition piece may be fabricated based on a constant slope of cross sectional area change between the aft end and the forward end.

Abstract: Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine.

Abstract: A combustor for a turbine engine is provided. The combustor includes a first liner and a second liner forming a combustion chamber with the first liner, the combustion chamber configured to receive an air-fuel mixture for combustion therein. The first liner defining an air admission hole for directing a first jet of pressurized air into the combustion chamber, and the air admission hole may have a non-circular shape. The combustor further includes an insert positioned within the air admission hole to guide the first jet through the air admission hole and into the combustion chamber.

Abstract: A turbomachine blade made of composite material including a fiber reinforcement obtained by three-dimensional weaving of yarns and densified by a matrix, comprises a first portion constituting an airfoil and a blade root. The first portion constitutes a single part with at least one second portion constituting a blade inner platform, said blade then lacking a blade anti-tilting wall, or a blade anti-tilting wall, said blade then lacking a blade inner platform. The first portion also constitutes a single part with at least one third portion constituting a blade outer platform spoiler plate, said blade then lacking a blade outer platform wiper plate, or a blade outer platform wiper plate, said blade then lacking a blade outer platform spoiler plate.

Abstract: The present invention provides a new and improved method of forming a turbine engine component having a plurality of airfoils disposed in an annular array between inner and outer shroud rings. The method includes forming inner and outer shroud ring patterns. The inner and outer shroud ring patterns may advantageously be formed by solid freeform fabrication techniques, such as stereolithography. An assembly fixture may be utilized to position the inner and outer shroud rings and an array of airfoils in a coaxial relationship. When the inner and outer shroud rings are to be assembled, axially inner and outer shroud ring patterns may be interconnected by providing relative rotation between annular sections of the shroud ring patterns. In addition, adhesive may be utilized to interconnect the sections of the shroud ring patterns.

Abstract: A sealing assembly for use with a rotary machine is described herein. The sealing assembly includes a stator shroud coupled to the casing. The stator shroud includes an inner surface that at least partially defines the cavity within the casing. At least one stator labyrinth tooth extends outwardly from the stator shroud inner surface towards a rotor assembly positioned within the casing. At least one protective member is coupled to the stator shroud upstream from the at least one stator labyrinth tooth to facilitate reducing a flow of combustion gas across the at least one stator labyrinth tooth.

Abstract: An apparatus for minimizing signature of an engine having an exhaust gas path and a fan gas path includes a casing, and a liner disposed within the casing. The exhaust gas path passes within the liner and the fan gas path passes between the liner and the casing. The casing is disconnected into an upstream portion and a downstream portion, each portion in registration with the liner, such that there are minimum pressure imbalances caused by the exhaust gas path, the fan gas path and ambient on the downstream portion.

Abstract: A fuel injector for a gas turbine engine may include a center body disposed about a longitudinal axis, and a premix barrel positioned radially outwardly from the center body to define an annular passageway between the center body and the premix barrel. The annular passageway may extend from an upstream end that is configured to be fluidly coupled to a compressor to a downstream end that is configured to be fluidly coupled to a combustor. The premix barrel may include a first portion at the upstream end and a second portion at the downstream end. The first portion may include a stainless steel material, and the second portion may include a nickel based superalloy material. The second portion may be coupled to the first portion by a laser clad coupling.

Abstract: A fuel nozzle for use with a turbine engine is provided. The fuel nozzle includes a housing coupled to a combustor liner defining a combustion chamber. The housing is at least partially positioned within an air plenum and comprises an endwall that at least partially defines the air plenum. The fuel nozzle includes a plurality of mixing tubes extending through the housing for channeling a fuel to the combustion chamber, a cooling fluid plenum at least partially defined within the housing by the housing endwall, and a plurality of apertures defined within the housing endwall for channeling a cooling fluid from the cooling fluid plenum to the air plenum.

Abstract: Systems, methods, and apparatus for linking machine stators are provided. A bracket may be utilized to link adjacent stators within a machine, for example, a turbine. The bracket may include a body portion, a first plurality of extensions, and a second plurality of extensions. The body portion may include a first edge operable to align with a base of a first stator and a second edge opposite the first edge and operable to align with a base of a second stator. The first plurality of extensions may extend from the first edge and may be operable to align with at least one corresponding groove on the base of the first stator. The second plurality of extensions may extend from the second edge and may be operable to align with at least one corresponding groove on the base of the second stator. When utilized to link the first stator and the second stator, the bracket may facilitate a reduction of vibrations in the first stator and the second stator.

Abstract: A method for machining a casing of an aircraft turboshaft engine, the casing including an upstream edge and a downstream edge. The method includes: positioning the casing in a chassis; rotating the chassis at a rotational milling speed; milling a downstream edge and/or upstream edge of the casing by a milling device, so as to correct axial length of the casing; and deburring the casing edge so as to remove at least one burr formed in a vicinity of the edge as a result of the milling, the milling device including a scraper tool which turn-machines the burr as the casing is rotated.

Abstract: Contaminant build-up and cooling airflow looses are reduced in a turbine airfoil by joining root and airfoil cooling air passages thereof with a transition passage.

Abstract: An example engine assembly includes a first attachment structure secured to an engine casing or an engine liner, and a second attachment structure secured to the other of the engine casing or the engine liner. A sliding member is held by the first attachment structure and is slideable relative to the first attachment structure between a first position and a second position. A pin structure moves with the sliding member between the first position and the second position. A link is pivotally connected to the second attachment structure and the pin structure.

Abstract: A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet at the first wall surface, an outlet at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally and laterally from the metering section, a second lobe adjacent the first lobe and diverging longitudinally from the metering section, a third lobe adjacent the second lobe and diverging longitudinally and laterally from the metering section, and a transition region having an end adjacent the outlet and a portion that extends between the lobes and the outlet. The first and third lobes each include a curved outer portion.

Abstract: Disclosed is a fuel nozzle end cover, a fuel nozzle and a process of fabricating a fuel nozzle end cover. The fuel nozzle end cover includes a base material and one or more features extending from the base material into a cavity of the fuel nozzle end cover. The one or more features are secured to the base material by a process selected from the group consisting of beam welding, friction welding, gas tungsten arc welding, gas metal arc welding, and combinations thereof. The fuel nozzle includes a fuel nozzle insert and a fuel nozzle end cover. The process of fabricating the fuel nozzle end cover includes providing a base material and securing one or more features to the base material by a welding process selected from the group consisting of beam welding, friction welding, gas tungsten arc welding, gas metal arc welding, and combinations thereof.

Abstract: A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section.

Abstract: A separate propulsion unit incorporating a free turbine and a fan receives gases from a plurality of core engines. The core engines each include a compressor, a turbine and a combustion section. The core engines in combination pass gases across the free turbine. A method is also disclosed.

Abstract: A gas turbine engine component includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface and a diffusing section in communication with the inlet and extending to the outlet. The diffusing section includes a plurality of crenellation features that encourage lateral spreading of cooling air flowing through the cooling hole.

Abstract: A gas turbine engine component includes a wall having first and second wall surfaces and first and second cooling holes extending through the wall. The first and second cooling holes each include an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. Each diffusing section includes first and second lobes, each lobe diverging longitudinally and laterally from the metering section. The outlets of each cooling hole include first and second lateral ends and a trailing edge. One of the lateral ends of the outlet of the first cooling hole and one of the lateral ends of the outlet of the second cooling hole meet upstream of the trailing edge of the first cooling hole and the trailing edge of the second cooling hole.

Abstract: The invention relates to a method for manufacturing an inexpensive micro gas turbine 1. According to this method, an existing compressor-turbine unit is divided into two separate parts being a compressor 3 and a turbine 5. The compressor and turbine shafts are joined. Two existing bearing units (11 and 17) are taken, wherein one of them may belong to the existing compressor-turbine unit. The rotor 25 of an existing generator 7 is mounted on the joined shaft. A generator housing 21 is manufactured and connected to the bearing units. The stator 23 of an existing generator 7 is mounted into the generator housing. Energy input into the working cycle of the gas turbine can be implemented by adding either an internal burner or external burner with a heat exchanger. By using inexpensive and often mass produced off-the-shelf components, a cost effective micro gas turbine can be derived.

Abstract: A turbofan engine has a fan having a circumferential array of fan blades. A fan case encircles the fan. There is at least one compressor section, a combustor, and at least one turbine section. The fan case comprises a composite structural member and a metallic member encircled by the composite structural member. The metallic member is mounted to the composite structural member to permit differential thermal expansion proximate the blades.