Abstract: A gear support assembly for a turbine engine includes an epicyclic gear arrangement and a first tapered bearing and a second tapered bearing spaced apart from the first tapered bearing. The first tapered bearing and the second tapered bearing are arranged axially forward of the epicyclic gear arrangement and support the epicyclic gear arrangement.

Abstract: A system, one embodiment, includes an annular seal configured to seal a turbine fuel nozzle to a turbine combustor, wherein the annular seal includes a first annular seal portion having an inner annular surface, a first base portion, and a first peripheral portion opposite from the first base portion. The embodiment also includes the first annular seal portion configured to couple to a second annular seal portion having an outer annular surface, a second base portion, and a second peripheral portion opposite from the second base portion. Further, the embodiment includes the first and second peripheral portions sealable by a weld.

Abstract: A combustor heat shield sealing arrangement comprises a sealing rail extending from the combustor liner shell at the exit of the combustor for sealing engagement with a rail-less downstream end portion of the combustor heat shield. The sealing rail is offset relative to the downstream vane passage. Doing so may minimize the combustor/vane waterfall and, thus, minimize the horseshoe vortex effect at the leading edge of the turbine vanes.

Abstract: A device for guiding an element in an orifice in a wall of a turbomachine combustion chamber is disclosed. The device includes a coaxial ring and bushing mounted one inside the other. The bushing includes two coaxial annular parts that are fastened to each other by welding and that define a groove for guiding a rim of the ring. Welding zones between the parts of the bushing are situated away from the path of the rim of the ring when moved along a privileged transverse direction of ring movement relative to the bushing when the combustion chamber is in operation.

Abstract: A wall element for a combustion space, the wall element comprising a ceramic body, a support within the body, and attachment means extending from the support and protruding from the body for securing the ceramic body to a combustion space wall.

Abstract: Sail wing aircraft which includes a wing (6) and at least one propulsion engine (8). It includes an upper beam (22) which is firmly fixed at its front end to a first frame (12) located on an air inlet (14) of the propulsion engine and which is in addition firmly fixed at its median part to a second frame (16) located to the rear of the first frame. The sail wing aircraft includes in addition a pylon (26) for attachment of the engine onto the fuselage, where the engine is fixed to the pylon (26).

Abstract: A transition duct used in a gas turbine engine to direct the hot gas flow from the combustor into the turbine section of the gas turbine engine. The transition duct includes a plurality of guide vane integral with the duct. The transition duct includes a circular shaped inlet end for connection to a can combustor and a rectangular and arched shaped outlet end for connection to a first stage turbine section. the guide vanes extend within the flow path between inner and outer projections each having a curved opening in the shape of the airfoil each airfoil includes inner and outer airfoil ends with retainer slots formed between the airfoil ends and the duct projections that form shear pin retainer slots. Shear pin retainers are secured within the slots to secure the guide vanes to the duct in a thermally uncoupled manner to reduce thermal stresses. The guide vanes can be made from a single crystal material for higher gas flow temperatures.

Abstract: A bracket for a combustor, and a method for mounting a transition piece in a combustor, are disclosed. The combustor has an impingement sleeve at least partially surrounding a transition piece and an outer casing at least partially surrounding the impingement sleeve. The bracket is mounted to the transition piece and connected to the outer casing. The method includes mounting a bracket to the transition piece, extending the bracket through an impingement sleeve, the impingement sleeve at least partially surrounding the transition piece, and connecting the bracket to an outer casing, the outer casing at least partially surrounding the impingement sleeve.

Abstract: A gas turbine engine has an engine casing component, such as a mid turbine frame system, which may includes an outer case surrounding a spoke or strut casing. Load transfer spokes of the spoke casing extend between the outer case and an inner case of the spoke casing. A bearing housing supported by the engine casing may include a fuse arrangement for isolating bearing seizure loads from the engine casing. A seal arrangement may be provided to centralize the rotors after the fuse fails. The mid turbine frame may be provided to have a desired flexibility through the configuration of the casing and bearings in design.

Abstract: A system is provided for mounting fuel injectors to a gas turbine engine. The system includes an engine casing having an aperture formed therein. The system further includes a plurality of fuel injectors having respective flanges for mounting the fuel injectors to the casing at the aperture so that the fuel injectors extend side-by-side into the engine. The flanges are dismountably sealed to an inner side of the casing. The aperture and the flanges are configured so that, when dismounted, each fuel injector can be rotated into an orientation relative to the aperture which allows the respective flange to pass though the aperture and the fuel injector to be withdrawn from the casing. The flanges have respective sealing formations which engage with their neighbouring flanges when the fuel injectors are mounted to the casing to close off the aperture and to form seals between the flanges.

Abstract: A power turbine receiving cooling fluid from compressor apparatus is provided. The turbine comprises: fixed housing structure; a support shaft rotatably supported in the housing structure; at least one disk overhung on the support shaft and coupled to the support shaft so as to rotate with the support shaft; and cooling fluid structure. The at least one disk is formed from a low-alloy-content steel. The cooling fluid supply structure extends through the support shaft and the at least one disk so as to supply cooling fluid to the at least one disk.

Abstract: Undue flexing and cracking of an internal exhaust shield of a turbine is eliminated by isolating thermally-induced movement in the shield from that of strut apparatus passing therethrough to a surrounding outer exhaust casing. Methods and apparatus are disclosed.

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.

Abstract: A dirt separator may be retrofit into existing gas turbine engines since it is formed of circumferentially separate pieces which may be assembled together in the gas turbine engine.

Abstract: A pylon for suspending an engine beneath an aircraft wing, capable of being attached by one end to a casing of the engine and by another end to the wing is disclosed. The pylon includes at least one articulation actuated by a actuator making it possible to change the height position of the engine on the ground and in flight.

Abstract: A condition measurement apparatus is provided and includes a gas turbine engine combustor having an end cover, a liner defining a liner interior and a fuel nozzle communicative with the liner interior, the end cover being formed to separate a cold side thereof, which is a relatively low temperature environment, from a hot side thereof, which is a relatively high temperature environment in which the liner and the fuel nozzle are disposed, the combustor being formed to define a fuel flow path extending through piping disposed at the cold side of the end cover by which fuel is deliverable to the fuel nozzle, and a condition sensing device operably mounted on the piping.

Abstract: A combination including a gas turbine engine extending along an axis is disclosed herein. The gas turbine engine includes an annular combustor having a combustor liner. The combination also includes a plurality of projections extending from the combustor liner and spaced from one another circumferentially about the axis. The combination also includes a free-standing ring disposed about the combustor liner and positioned adjacent to the plurality of projections along the axis. The plurality of projections engage the free-standing ring and circumferentially support the combustor liner while allowing relative radial displacement between the combustor liner and the free-standing ring. The combination also includes a plurality of pins each being integrally-formed with one of the plurality of projections. Each of the plurality of pins extends along the axis and is received in one of a plurality of slots formed in the free-standing ring.

Abstract: The present invention relates to a nacelle for a turbojet, comprising an air inlet structure capable of channelling a flow of air to a turbojet fan and a middle structure (5) comprising a casing (9) designed to encircle said fan, with the air inlet structure attached to it, said structure having a peripheral inside surface at least partly equipped with a sound-attenuating structure (13) extending without geometrical interruption around at least part of the casing, said structure being characterized in that a gap (14) is provided between the sound-attenuating structure and the casing.

Abstract: A support between a transition piece and an impingement sleeve in a combustor is disclosed. The support includes a resilient portion, the resilient portion configured to provide dampening between the transition piece and the impingement sleeve. The support further includes a mount portion configured for mounting the support to one of the transition piece or the impingement sleeve. The support further includes a contact portion configured for contacting the other of the transition piece or the impingement sleeve.

Abstract: An assembly for gas turbine engine includes a cross-over housing, an inter-shaft seal housing, and a linkage. The inter-shaft seal housing is disposed within the cross-over housing. The linkage is connected to the inter-shaft seal housing and extends through a slot in the cross-over housing. The linkage is removable from both the inner radial housing and the outer radial housing.

Abstract: An engine unit for an aircraft including a turboprop engine and its device for mounting on to a wing surface. The device includes a rigid structure and a mechanism for fastening the turboprop engine on to this structure. The fastening mechanism includes six mutually independent hydraulic systems, each one exclusively dedicated to the transfer, to the rigid structure, of forces exerted respectively according to one of the six degrees of freedom of movement. Each hydraulic system includes at least one hydraulic jack with a piston attached to one of the two elements, i.e. either the turboprop engine or the rigid structure, together with a cylinder housing the piston and attached to the other of the two elements.

Abstract: According to one aspect of the invention, a shim for sealing two adjacent turbine transition pieces is disclosed. The shim includes a circumferential member that includes a first lateral flange and a second lateral flange. Further, the first and second lateral flanges each comprise a tab configured to mate to a first surface plane and the first and second lateral flanges are configured to mate to a second surface plane, wherein the first and second surface planes are substantially parallel. In addition, the shim includes a first flange extending substantially perpendicular from the circumferential member.

Abstract: A bleed valve module for mounting to a support structure within an aircraft includes a pipe, a plurality of bleed valves connected to the pipe, a plurality of pneumatic actuators connected to the plurality of bleed valves, and a vibration isolating element. The pipe contains a flow of bleed air and includes a plurality of inlets and an outlet. The plurality of bleed valves controls the flow of bleed air in the pipe. The plurality of pneumatic actuators actuate the plurality of bleed valves. The vibration isolating element connects the bleed valve module to the support structure within the aircraft and isolates the bleed valve module from vibration.

Abstract: A system, including, a turbine combustor, including, a first wall disposed about a flow path of hot combustion gases, a second wall disposed about the first wall, and a damping system disposed between the first and second walls, wherein the damping system is configured to dampen vibration, and the damping system is tuned to dynamic drivers in the turbine combustor.

Abstract: Gas turbine engines and related systems involving offset hub struts are provided. In this regard, a representative bearing assembly for a gas turbine engine includes: a bearing operative to support a rotatable shaft; an annular hub positioned about the bearing; and an annular array of struts extending radially outwardly from the hub, at least two of the struts being positioned in different planes, the planes being oriented transversely with respect to the rotatable shaft. The method of assembling the struts by mounting first and second pluralities of struts, each plurality in a common plane, between the hub and the turbine exhaust case, with the second plurality of struts longitudinally offset from the first plurality of struts.

Abstract: A suspension of a multi-flow turbojet engine provided with an intermediate casing and an exhaust casing from a pylon that can be attached to the structure of an aircraft is disclosed. The suspension includes a forward attachment device between the hub of the intermediate casing and the pylon, a rear attachment device between the exhaust casing and the pylon, and a connection device rigidly connecting the intermediate casing to the pylon. The rear attachment device includes an actuator for compensating for the variations in diameter of the exhaust casing so as to keep the axis of the exhaust casing coaxial with the axis of the intermediate casing through the various phases of flight of the aircraft.

Abstract: A device for attaching an aircraft engine including a thrust force take-up device that includes a main fitting to be fixed to an engine mounting beam fixed to a box structure, an additional fitting fixed to the main fitting, and two load reacting link rods connected to the main fitting by a balance beam, each of the link rods being connected to the balance beam by a first longitudinal end by a first mechanical connection that has at least one axis of rotation, each of the link rods being connected to the additional fitting by a second longitudinal end by a second mechanical connection that has at least one axis of rotation, the second mechanical connection having play, the first and second mechanical connections having at least one axis of rotation in common.

Abstract: In a gas turbine engine combustor and method for delivering purge gas thereto, a ferrule is generally coupled to the housing and moveable relative thereto. The ferrule has a primary opening through which a combustor component extends into a combustion chamber of the combustor. The ferrule further has a plurality of purge gas openings separate from and in transversely spaced relationship with the primary opening to allow purge gas to flow through the ferrule. In a first position of the ferrule at least one purge gas opening is blocked against the flow of purge gas therethrough and at least one other purge gas opening is unblocked against purge gas flow. In a second position of the ferrule at least one of the blocked purge gas openings of the first ferrule position is unblocked to permit the flow of purge gas therethrough.

Abstract: An annular shroud surrounding tip ends of turbine blades is fitted on an inner peripheral surface of an open end of a turbine case. In order to prevent the shroud from being dropped off in an axial direction by a retaining part provided on a retaining ring fitted on an outer peripheral surface of the open end of the turbine case, an annular groove formed in the outer peripheral surface of the turbine case and an annular groove formed in an inner peripheral surface of the retaining ring are opposed to each other, and a connection wire is inserted so as to straddle the two annular grooves, thereby connecting the retaining ring to the turbine case. Accordingly, it is possible to easily perform attachment/detachment of the retaining ring with respect to the turbine case only by inserting the connection wire into or pulling it out from the two annular grooves, thereby facilitating maintenance of the shroud and the turbine blades.

Abstract: An impingement shield assembly for a turbine includes a first impingement shield portion and a second impingement shield joined to the first impingement shield portion. The assembly also includes a first connection portion formed on the first impingement shield portion, a second connection portion formed on the first impingement shield portion and first and second wedge weld portions that mate with and hold the first and a second connection portions in a fixed relationship to one another.

Abstract: A system includes a gas turbine engine, a nozzle structure located adjacent to the gas turbine engine, a flexible annular seal linking the gas turbine engine and the nozzle structure, the flexible annular seal having a plurality of corrugate sections arranged about a centerline, and a connection device secured to the gas turbine engine and to the nozzle structure. The connection device is adjustable to select a length of the flexible annular seal measured along the centerline.

Abstract: The positioning arrangement comprises at least one pair of complimentary rounded recesses, each recess being provided on the mating face of the corresponding component. One ball is positioned between each pair of complementary recesses.

Abstract: A mid-turbine frame connected to at least one mount of a gas turbine engine transfers a first load from a first bearing and a second load from a second bearing to the mount. The mid-turbine frame includes a single point load shell structure and a plurality of struts. The single point load shell structure combines the first load and the second load into a combined load. The plurality of struts is connected to the single point load structure and transfers the combined load from the single point load shell structure to the mount.

Abstract: A gas turbine engine is provided having an annular duct having an inner duct wall and an outer duct wall. At least one strut extends between the inner duct wall and the outer duct wall. The strut has an aerodynamic shape. At least one vane located circumferentially adjacent the strut and axially behind a leading edge of the strut, and an aerodynamic deflector located circumferentially between each strut and each vane. The aerodynamic deflector may positioned axially upstream of a flow separation point which would otherwise be caused by the strut.

Abstract: A single point load structure transfers a first load from a first bearing cone and a second load from a second bearing cone of a gas turbine engine to a plurality of struts. The single point load structure includes a stem, a branch connected to the stem and a torque box connected to the plurality of struts for absorbing the first and second loads from the stem and the branch. The stem has a concave surface that opens in a radially outward direction with respect to a rotational axis of the gas turbine engine.

Abstract: An extension of an intermediate casing for an aircraft jet engine, intended to project towards the downstream end from an external shell of said intermediate casing is disclosed. This extension includes an annular connecting downstream end forming an annular groove open radially towards the outside, intended to receive nacelle covers. The annular connecting downstream end is produced by a plurality of angular sectors.

Abstract: A mounting system supporting and positioning an internal fuel manifold of a gas turbine engine includes at least a heat shield surrounding a fuel inlet of the fuel manifold, the heat shield bearing at least a portion of a load to support the fuel manifold.

Abstract: An igniter assembly (10) of a gas turbine (12) is presented, including an igniter (14) disposed within an igniter housing (16). The igniter is extendable from the igniter housing through an opening (18) in a combustion liner (19) to an extended position, and is retractable from the extended position back through the opening to a retracted position (22). The igniter assembly further includes a compressible assembly (24) disposed between a base (26) of the igniter housing and the combustion liner to form a sealed interface (15) with a perimeter (28) of the opening (18). The compressible assembly is configured to restrict an air flow from passing between the igniter housing base and the perimeter of the opening. The compressible assembly is variable in length to accommodate a respective variation in a separation between the igniter housing base (26) and the opening (18).

Abstract: A combustor liner may include an annular inner liner and an annular outer liner with a plurality of air holes thereon. The outer liner may be positioned circumferentially around the inner liner such that an annular cooling space is defined between the inner and the outer liner. The combustor liner may also include at least one resonator coupled to the outer liner such that a base of the resonator is separated from the outer liner to form a gap with an external surface of the outer liner. The combustor liner may also include a throat extending from the base of the resonator penetrating the inner liner and the outer liner. The combustor liner may further include a grommet assembly that allows for relative thermal expansion between the inner liner and the outer liner proximate the throat.

Abstract: An assembly for a gas turbine engine includes a combustor and a vane assembly disposed downstream thereof. A portion of an outer platform of the vane assembly defines an axial sliding joint connection with the combustor, and includes a plurality of depressions located in an outer circumferential surface thereof opposite the combustor. The depressions are disposed in regions of expected higher thermal growth about the circumference of the outer platform such that thermal growth of the entire outer platform is substantially uniform circumferentially therearound.

Abstract: According to one aspect of the invention, a shim for sealing two adjacent turbine transition pieces is disclosed. The shim includes a circumferential member that includes a first lateral flange and a second lateral flange. Further, the first and second lateral flanges each comprise a tab configured to mate to a first surface plane and the first and second lateral flanges are configured to mate to a second surface plane, wherein the first and second surface planes are substantially parallel. In addition, the shim includes a first flange extending substantially perpendicular from the circumferential member.

Abstract: An engine mounting apparatus that includes thrust struts, a balance yoke, and primary and secondary mount blocks. The thrust struts are connected to the balance yoke and the primary mount block has at least one thrust load connection connectable to the balance yoke for transferring thrust loads from an engine during normal operation of the engine mounting apparatus. The engine mounting apparatus is characterized in that the secondary mount block is connected to the balance yoke via at least one thrust load catcher connection, the at least one thrust load catcher connection being arranged to be unloaded during normal operation of the engine mounting arrangement and to transfer the thrust loads from the engine in the event of a failure condition of the primary mount block.

Abstract: An assembled annular combustion chamber comprises an annular inner wall and an annular outer wall made of ceramic matrix composite material together with a chamber end wall connected to the inner and outer walls and provided with orifices for receiving injectors. Elastically-deformable link parts connect the inner wall and the outer wall of the chamber to inner and outer casings that are made of metal. The assembly formed by the inner wall, the outer wall, and the combustion chamber end wall is subdivided circumferentially into adjacent chamber sectors, each sector being made as a single piece of ceramic composite material and comprising an inner wall sector, an outer wall sector, and a chamber end wall sector. The link parts connect the inner metal casing and the outer metal casing respectively to each inner wall sector of the combustion chamber and to each outer wall sector of the chamber. The chamber end wall sectors are in contact with a one-piece ring to which they are connected.

Abstract: Methods and apparatus for a mounting assembly for a liner of a gas turbine engine combustor are provided. The combustor includes a combustor liner and a radially outer annular flow sleeve. The mounting assembly includes an inner ring surrounding a radially outer surface of the liner and including a plurality of axially extending fingers. The mounting assembly also includes a radially outer ring coupled to the inner ring through a plurality of spacers that extend radially from a radially outer surface of the inner ring to the outer ring.

Abstract: A turbomachine fan duct is disclosed. The fan duct includes two coaxial cylindrical walls, these being an internal wall and an external wall, respectively. The internal wall includes a framework to which panels are removably attached, and the external wall includes a single support structure with openings closed by removable panels. The openings in the external wall have dimensions that allow the panels of the internal wall to pass and allow these panels to be fitted onto and removed from the framework of the internal wall.

Abstract: A combustion gas measurement apparatus mounted in a gas turbine including: a tunable laser generating a radiation beam passing through a combustion gas path; a controller tuning the laser to emit radiation having at least a first selected wavelength and a second selected wavelength which both correspond to temperature-dependent transitions of a combustion species of the gas, wherein the first selected wavelength and the second selected wavelength are not near absorption peaks of neighboring wavelengths; a detector sensing the radiation beam passing through the combustion gas and generating an absorption signal indicative of an absorption of the beam by the combustion gas at each of the first wavelength and the second wavelength, and a processor executing a program stored on a non-transitory storage medium determining a combustion gas temperature based on a ratio of the adoption signals for the first wavelength and the second wavelength.

Abstract: A gas turbine system includes a combustion chamber (2), a turbine (3), a radially inward and/or radially outward axial gap (4; 16, 17) between the combustion chamber (2) and the turbine (3), at which gap inner and/or outer combustion chamber walls (7, 8) and inner and/or outer turbine walls (11, 12) end, and a cooling gas supply (5), which via the gap (4; 16, 17) introduces a cooling gas into the turbine gas path (13) and/or into the combustion chamber gas path (9). An end section (21, 22, 23, 24) of the respective turbine wall (11, 12) and/or of the respective chamber wall (7, 8) adjoining the gap (4; 16, 17) is radially inwardly and/or radially outwardly, alternatingly formed.

Abstract: A hot gas path system for a gas turbine including a stepped inlet ring at an intersection between a downstream end of a combustor basket and an upstream end of a transition is disclosed. The heat shield may be positioned downstream from a combustor basket and proximate to an intersection between the collar and the axially extending cylindrical wall. The stepped inlet ring may be coupled to the transition section and may extend upstream from the transition section. An upstream end of the stepped inlet ring may be positioned radially outward from the combustor basket such that at least a portion of the stepped inlet ring overlaps a portion of the combustor basket. A spring clip may be positioned between the combustor basket and the stepped inlet ring such that the spring clip seals at least a portion of a gap between the combustor basket and the stepped inlet ring.

Abstract: Systems and methods provide for reducing motion which causes wear in a combustor. The method includes attaching a first end of a combustor liner to a retainer which encircles the first end of the combustor liner; attaching the retainer to a combustor casing, where the retainer acts as a spring between the combustor liner and the combustor casing; reducing a relative motion between the combustor casing and the combustor liner along a longitudinal axis and a circumferential axis of the combustor liner by the retainer; and locating the combustor liner substantially concentrically within the combustor casing.

Abstract: A combustion chamber of a combustion system has a combustion space, a support structure, a support element, and a heat shield. The heat shield has at least two segments, and each segment includes a liner element facing the combustion space and has an edge region, a gap communicating with the combustion space being formed between edge regions of adjacent segments, and a retaining device. The retaining device fixes the respective liner element on the support structure via the support element and forms a flange region that fits over the edge region of the respective liner element. The retaining device forms a first cooling passage with the support element and has at least one through-opening in the flange region. A cooling gas flows through the through.opening from the first cooling passage to the edge region.