Abstract: A fuel assembly for a gas turbine engine includes a fuel supply tube, a fuel port, a fuel manifold, and a fuel manifold adapter. The fuel supply tube is configured to convey a fuel. The fuel port is fluidly coupled to the fuel supply tube and configured to receive the fuel from the fuel supply tube. The fuel manifold includes a fuel inlet and a plurality of fuel outlets. The fuel inlet is fluidly coupled to the fuel port and configured to receive the fuel from the fuel port. The fuel manifold adapter includes a first mount portion and a second mount portion. The first mount portion is connected to the fuel port. The first mount portion is moveable relative to the second mount portion.

Abstract: An aircraft propulsion unit includes a drive unit with a static part and a rotary part which rotates a fan situated downstream from the drive unit, an assembly of fixed blades situated downstream from the fan, and a nacelle in which the fan and the assembly of fixed blades are accommodated. The propulsion unit also includes an assembly of at least two coaxial shafts, wherein a fan shaft connects the fan to the rotary part, and a stator blading shaft connecting the assembly of fixed blades to the static part extends concentrically, and for at least part of its length in the interior of the fan shaft. This rigid and compact configuration limits the variations of distance between the end of the fan blades and a fan housing situated in the inner duct of the nacelle.

Abstract: Integrated combustor nozzles and turbomachines are provided. An integrated combustor nozzle includes a cooling insert having a flange. The cooling insert further includes a first wall and a second wall that each define respective passages therein. The first wall and the second wall each extend from an open end defined within the flange to a closed end. The first wall and the second wall each include an impingement side spaced apart from a solid side. Each impingement side defines a plurality of impingement apertures configured to direct coolant from the respective passage towards a side wall of the combustion liner. The cooling inserts further includes a collection passageway that is defined between the solid sides of the first wall and the second wall.

Abstract: A turbomachine including a stator, a rotor mounted on a shaft, a first bearing and a second bearing downstream of the first bearing in a direction of air circulation at the inlet of the turbomachine, the first bearing forming an axial abutment linked to the stator, the second bearing including an inner ring secured to the shaft and an outer ring mounted on an arm secured to the stator, the second bearing including rolling elements in axial support on the inner ring and on the outer ring. The outer ring includes a radially extending ring portion. The turbomachine also includes an axial force application device mounted between the arm and the ring portion.

Abstract: A combustor for use in a gas turbine engine, the combustor enclosing a combustion chamber having a combustion area. The combustor includes: a shell having a kink; and a kinked heat shield panel in facing spaced relationship with the shell, the kinked heat shield panel including a kink located proximate the kink in the shell, wherein the kinked heat shield panel further includes a first surface, a second surface opposite the first surface, and a mounting stud located proximate the kink of the kinked heat shield panel and extending away from the second surface.

Abstract: A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Impingement tubes direct cooling air onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine.

Abstract: The gas turbine engine fuel manifold assembly includes a fuel manifold, and fuel nozzles secured to the fuel manifold and in fluid communication therewith. A fuel inlet connector has a body secured to the fuel manifold, and is in-line therewith. The fuel inlet connector has at least one connector outlet fluidly connected to the fuel manifold. The at least one connector outlet defines a fuel flow direction having a main component in a circumferential direction relative to the axis.

Abstract: A method of constructing a gas turbine system may include placing a foundation interface of a gas turbine mannequin on a foundation, aligning the gas turbine mannequin with a load and connecting the gas turbine mannequin to the load. The method may include constructing a pipeline and connecting the pipeline to the gas turbine mannequin. The method may include placing a foundation interface of a gas turbine mannequin on a foundation, aligning the gas turbine mannequin with an exhaust structure and connecting the gas turbine mannequin to the exhaust structure. The method also includes removing the gas turbine mannequin and installing a gas turbine.

Abstract: An internal fuel manifold for a turbomachine can include an outer ring defining one or more fuel supply channels configured to allow fuel flow around at least a portion of a circumference of the outer ring, and a plurality of branches extending radially inward from the outer ring. The branches include a high aspect ratio shape. The plurality of branches can include one or more branch fuel channels in fluid communication with at least one of the one or more fuel supply channels. The manifold includes one or more injector connectors extending axially aft from each branch, each injector connector in fluid communication with a branch fuel channel and configured to connect to a fuel injector. The manifold can include an inner ring extending from the plurality of branches. The inner ring and the outer ring are configured to receive a fuel injector and combustor assembly therebetween and to allow retaining of the fuel injector assembly and combustor assembly to the internal fuel manifold.

Abstract: A gas turbine engine may include a fuel manifold extending circumferentially around a diffuser case of the gas turbine engine. The fuel manifold may include a fuel supply inlet interface for receiving fuel into the fuel manifold and a plurality of fuel delivery outlet interfaces for delivering fuel to a combustor of the gas turbine engine. The gas turbine engine may also include a valve coupled to the fuel manifold. The valve may be configured to control fuel distribution in the fuel manifold. The valve may be disposed between two fuel delivery outlet interfaces of the plurality of fuel delivery outlet interfaces. The valve may be configured to at least decrease fuel flow to one of the plurality of fuel delivery outlet interfaces. The valve may be configured to at least decrease fuel flow to half of the plurality of fuel delivery outlet interfaces.

Abstract: Gas turbine engine systems involving integrated fluid conduits are provided. In this regard, a representative engine casing for a gas turbine engine includes: a casing wall; and a fluid conduit integrated into the casing wall.

Abstract: A liquid fuel injection spoke for a gas turbine engine. The liquid fuel injection spoke may include an inlet end and an outlet end. The liquid fuel injection spoke may further include a head located at the inlet end and a stem extending along a longitudinal axis from the head to the outlet end. The stem may define a fluid passageway therein. The head may include a cavity with an inner wall. The inner wall may include an annular protrusion.

Abstract: An engine mounting arrangement is provided for supporting a gas turbine engine. The mounting arrangement comprises a module adaptor mounted to an overhead or a ground based carrier for rotation about a substantially horizontal axis. The module adaptor extends horizontally from a base to an axially-opposed distal end having a first mounting interface configured for attachment to a module of the engine. The module adaptor holds the engine by the module to which it is attached and with the other engine modules projecting in a cantilever fashion from the supported module.

Abstract: Embodiments of the present disclosure provide a turbine combustor having a primary fuel injection system, a first wall portion disposed about a primary combustion zone downstream from the primary fuel injection system, and a second wall portion disposed downstream from the first wall portion. The turbine combustor also has a secondary fuel injection system disposed between the first wall portion and the second wall portion, where the secondary fuel injection system is removable form the first and second wall portions.

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 system for packaging a gas turbine includes an enclosure peripherally surrounding and extending along a length of the gas turbine. The enclosure comprises first and second ends with a passage between the enclosure and the gas turbine. The system further includes means for connecting the first and second ends to an air filtration system and an exhaust system, respectively. A plurality of supports are connected to and extend radially inward from the enclosure with means for connecting the plurality of supports to the gas turbine. A method for packaging a gas turbine includes connecting the gas turbine to a plurality of supports and connecting the plurality of supports to an enclosure peripherally surrounding and extending along a length of the gas turbine. The method further includes connecting first and second ends to an air filtration system and an exhaust system, respectively.

Abstract: The invention relates to a nacelle rear assembly for a turbojet engine, that comprises: a hood, an inner structure in the shape of a sheath in which at least the rear portion (15) is of the O-duct type and is capable of axial sliding between an operational position in which it covers the gas generator (3) of said turbojet engine and defines an annular cold-air jet with said hood, and a maintenance position downstream from said operational position. The downstream portion (15) is conformed so as to be capable of sliding without blocking relative to said gas generator (3), and the inner structure includes an upstream portion (13) that can be separated from the downstream portion (15) and that includes at least two doors (13a, 13b) capable of opening towards the outside.

Abstract: One embodiment of the present application includes a hot section component of a gas turbine engine having a covering. The covering includes a protrusion and is attached to the hot section component though a flexible retainer. In one form the covering is made from ceramic matrix composite. The flexible retainer has a closed position and an open position. The retainer secures the protrusion to the hot section component when it engages part of the protrusion when in the closed position.

Abstract: An alignment and measuring tool for turbomachine combustor includes a support member configured and disposed to operatively connect to a combustor casing mounting element having an opening, and an indexing element connected to the support member. The indexing element includes an opening provided with a first recess zone and a second recess zone. A ferrule including an opening is configured and disposed to be positioned across a combustion liner opening. A cap member is supported upon the support member and arranged within the opening of the indexing element. A rod is arranged in the cap member and extends from a first end to a second end through an intermediate portion. The rod includes an adjustable cone member slidingly mounted upon the intermediate portion. The adjustable cones is configured and disposed to nest in the combustor liner opening.

Abstract: A mounting assembly 2 for mounting a gearbox 10 to a gas turbine engine, comprising a mount ring 4 and a spigot pin 8 having a flange 22 at one end. The spigot pin 8 extends radially outwardly through a bore 14 in the mount ring 4 for engagement with a gearbox 10 and the flange 22 locates against a radially inner surface of the mount ring 4. The bore 14 supports the spigot pin 8 and dissipates forces exerted on the spigot pin 8, for example forces generated during a fan blade-off event.

Abstract: A split fuel control module for a gas turbine engine and method of installation. The split fuel control module includes a first frame unit, a second frame unit, a segmented fuel path, and a distributed fuel controller. The first frame unit and the second frame unit are joined together at a frame unit interface. The segmented fuel path includes an upstream fuel interface fixed to the first frame unit and a downstream fuel interface fixed to the second frame unit and detachably coupled to the upstream fuel interface at the frame unit interface. A first portion of the distributed fuel controller is fixed to the first frame unit, and a second portion of the distributed fuel controller is fixed to the second frame unit.

Abstract: A system includes a gas turbine combustor, which includes a combustion liner disposed about a combustion region, a flow sleeve disposed about the combustion liner, an air passage between the combustion liner and the flow sleeve, and an aerodynamic mounting assembly disposed in the air passage. The aerodynamic mounting assembly is configured to retain the combustion liner within the flow sleeve. The aerodynamic mounting assembly includes a flow sleeve mount coupled to the flow sleeve and a liner stop coupled to the combustion liner. The flow sleeve mount includes a first portion of an aerodynamic shape and the liner stop includes a second portion of the aerodynamic shape, which is configured to direct an airflow into a wake region downstream of the aerodynamic mounting assembly. The flow sleeve mount and the liner stop couple with one another to define the aerodynamic shape.

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 turbojet engine with a fan and a nacelle element forming the secondary duct downstream of the fan is provided. The turbojet engine includes an intermediate casing with an outer shroud to which the nacelle element is fixed by a connecting member. The connecting member includes a cylindrical element secured to the shroud while at the same time being capable of a rotational movement with respect to the axis of the engine and connected to the nacelle element by a bayonet-type fitting.

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: A gas turbine engine includes a fan including a plurality of fan blades rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a geared architecture arranged in a housing and driven by the turbine section for rotating the fan about the axis, and a support member that supports the geared architecture within the gas turbine engine. The support member includes an inner portion and an outer portion. One of the inner and outer portions is configured to be coupled to the geared architecture and the other of the inner and outer portions is configured to be coupled to the housing. A plurality of removal features each have at least one engaging surface to facilitate a pulling force on the support member in a direction parallel to the axis.

Abstract: A gas turbine engine adapter is provided. In one form the adapter is segmented and may be attached to a gas turbine engine support ring. The adapter can provide a channel that may receive part of a gas turbine engine component such as, but not limited to, a plug nozzle. The adapter can include a load bearing support to counteract axial forces experienced by the plug nozzle during operation of the gas turbine engine. A compressible seal can be provided between the plug nozzle and the adapter. A leaf seal can also be provided to seal between segments of the adapter. The plug nozzle can be discouraged from rotating relative to the adapter.

Abstract: The invention relates to a nacelle for a turbojet engine that includes an air intake structure capable of directing an air flow towards a fan of the turbojet engine, and a middle structure for surrounding said fan and to which is attached the air intake structure in order to provide aerodynamic continuity, wherein the air intake structure includes at least one inner panel attached to the middle structure and defining, with the latter, a nacelle fixed structure and at least one outer longitudinal panel removably attached to the fixed structure and including an air intake lip in order to define a removable air intake structure, wherein the removable air intake structure is provided with a locking device including at least one electric lock capable of interaction with a complementary retention structure.

Abstract: A novel and improved swirler assembly for use in a combustion liner of a gas turbine engine is disclosed. The swirler assembly is removable and sized to provide an improved assembly, repairability, and fit-up in the combustion liner. The swirler assembly includes a premix tube, premix swirler and a plurality of axially spaced mounting blocks for receiving removable fasteners for securing the swirler assembly to the combustor.

Abstract: A liner stop for retaining a flow sleeve relative to a combustion liner in a combustor is disclosed. The liner stop includes a female component defining a recess and a male component comprising a protrusion, the protrusion insertable in the female component. The liner stop further includes an insert disposable between the female component and the male component, the insert comprising a surface protrusion configured to interfere with one of the female component or the male component.

Abstract: A combustor liner for a turbine engine is disclosed herein. The combustor liner includes an inner liner surface operable to define at least part of a combustion chamber in a turbine engine. The inner liner surface extends along a portion of a chordal arc on a first side of the chordal arc. The combustor liner also includes a bearing surface operable to support a floating dome panel. At least part of the bearing surface is spaced from the chordal arc on a second side of the chordal arc opposite the first side.

Abstract: An aircraft nacelle that makes it possible to channel a stream of air in a longitudinal direction and that includes at least two panels (50, 50?) placed end to end in the longitudinal direction, each including a so-called aerodynamic surface (52, 52?) that is in contact with the air stream and an edge (54, 54?) that is in contact with the edge of the other panel, characterized in that at least one of the panels includes—at its edge (54, 54?)—at least one partition (58) that projects relative to the edge that extends in the direction of the edge of the other panel, in a manner that is secant to the aerodynamic surface and in a plane that contains the longitudinal direction, so as to limit the circulation of air between the edges of the panels.

Abstract: A gas turbine includes a gas turbine body in which a rotor is rotated with the energy of combustion gas produced by combustion of fuel to take out rotational energy from the rotor, an inlet casing that is attached to the gas turbine body and that guides the air to a compressor section compressing the air, an air intake chamber that is connected to the inlet casing and that guides the air taken in from the atmosphere to the gas turbine body, and a notch so formed in the air intake chamber as to cover the entire range where the rotor passes or exists when the rotor is moved.

Abstract: A splitter is disclosed that can be coupled with a splitter support and used within a diffuser of a gas turbine engine. The splitter includes apertures for receiving a portion of the splitter support. The splitter support includes support arms that are adapted to be slidingly received within the apertures of the splitter.

Abstract: A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A first shaft supports a low pressure compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A first bearing supports the first shaft relative to the inlet case. A second bearing supports a second shaft relative to the intermediate case. A low pressure compressor hub is mounted to the first shaft. The low pressure compressor hub extends to the low pressure compressor section between the first bearing and the second bearing.

Abstract: A heat shield element arrangement including a heat shield element for a heat shield arranged on a supporting structure is provided. On each of the two opposing sides running parallel to the installation grooves the heat shield element includes a continuous screw head opening which penetrates the cold side and the hot side of the heat shield element substantially vertically and through which the screw head of the respective screw is arranged to be accessible to the supporting structure, and a spring element may be arranged under the respective screw, which spring element may be extended along the hot side of the heat shield element and the parallel installation groove of the supporting structure. An outer end of the spring element is embodied as a clamping retaining hook which is provided for the purpose of engaging in the laterally recessed retaining groove of the heat shield element.

Abstract: An air intake structure is mounted upstream of a mid-structure of a nacelle for an aircraft engine. The air intake structure includes an external wall incorporating a lip, a partition defining a deicing compartment in the lip, a deicing manifold extending in the compartment, and a duct for supplying hot air to the deicing manifold. In particular, the external wall is mounted movably with respect to the mid-structure between a rear position and a front position, and the duct is connected in a fixed manner to the mid-structure. The partition is fixed in a sealed manner inside the lip, and the duct is connected to the manifold by a disconnectable seal which is situated in an immediate vicinity of the partition.

Abstract: A turbine system includes a compressor unit which compresses a fluid; a combustor unit which burns fuel with the compressed fluid; a turbine unit which is driven due to a combustion gas that is generated when the combustor unit burns the fuel; and a first device which receives power that is generated when the turbine unit is driven. The at least one compressor unit operates using at least the generated power, and is separately disposed from at least one of the turbine unit and the first device.

Abstract: A first member is retained to a second member to resist separation in a first direction. A joint comprises a first recess in the first member. The first member comprises at least a ceramic matrix composite (CMC) substrate. The joint comprises at least one ceramic key partially accommodated in the first recess and engaging the second member. The second member comprises at least a metallic substrate.

Abstract: A nacelle for a turban engine includes an air intake structure and a central structure. The central structure has a casing intended to surround a fan and an external structure. The air intake structure has at least one internal panel attached to the central structure via the casing and therewith forming a fixed structure and at least one longitudinal external panel attached removably to the fixed structure and incorporating an air intake lip. The external panel is moveable between an operating position and a maintenance position in which the external panel is separated from the external structure of the central section and the air intake lip is separated from the internal panel of the air intake structure.

Abstract: A gas turbine engine includes a compressor having first annular connection structure at a forward end portion of an air inlet housing of the compressor. The gas turbine engine may also include intake ducting having second annular connection structure adjacent a central passage of the intake ducting. The first annular connection structure and the second annular connection structure may form at least part of an interface between the air inlet housing and the intake ducting during operation of the gas turbine engine. The first and second annular connection structures may have shapes allowing the first annular connection structure to be moved through and beyond the second annular connection structure in a forward direction, the forward direction being opposite a direction of compressed air flow in the compressor during operation.

Abstract: A core engine for a gas turbine engine, the core engine having a rotational axis, a casing and further comprises a gearbox assembly, a radial drive arrangement, pipes and connections and a connection structure; the radial drive arrangement drivingly connects the core engine to the gearbox assembly; the pipes and connections extend from the gearbox assembly to the core engine and the gearbox assembly is radially spaced from the casing by connection structure. A method of assembling the gas turbine engine includes the steps of axially translating the core engine and the fan casing so that the gearbox is adjacent the fan casing and attaching the gearbox assembly to the fan casing.

Abstract: A bushing for assembling an exhaust duct liner with an exhaust duct in a gas turbine engine comprises a body, a bushing opening and a first tab. The body is insertable into a duct opening in an exhaust duct. The bushing opening extends through the body to receive a shaft of a liner fastener. The first tab protrudes from the body to extend into an interior of the duct to prevent rotation of a head of the liner fastener.

Abstract: A pressure-loss adjusting-member installation tool having pressure-loss adjusting-member grippers that are configured to freely grip or release a pressure-loss adjusting member which is arranged at a position between a lower nozzle arranged at one end of a fuel assembly having a relatively small pressure-loss when a primary coolant flows and a lower core support plate, and which can cause the primary coolant to pass therethrough in a state with a pressure-loss being increased more than that of when the primary coolant passes through only core support plate flow holes formed on the lower core support plate; and a holding unit that holds the pressure-loss adjusting-member grippers in a same positional relationship as a relative positional relationship between the fuel assemblies arranged in plural on the lower core support plate.

Abstract: A power system includes a gas turbine engine, which may include a gear box. The gear box may include a pipe connection configured to receive a removable pipe. The power system may also include a moveable support configured to fasten to the pipe connection to at least partially support the gear box.

Abstract: A turbomachine combustion chamber including a substantially L-shaped fuel injector and a device for mounting and fastening the injector in an orifice of the casing of the chamber, is disclosed. The device includes an outer collar carried by the injector and a removable annular spacer surrounding the injector and designed to be interposed between the collar of the injector and the casing.

Abstract: In an SEV combustor and a method for reducing emissions in an SEV combustor of a sequential combustion gas turbine, an air/fuel mixture is combusted in a first burner and the hot gases are subsequently introduced into the SEV combustor (1) for further combustion. The SEV combustor (1) includes a chamber having a chamber wall (5) defining a mixing portion (8), for mixing the hot gases with a fuel, and a combustion region (9), at least one inlet (2) for introducing the hot gases into the mixing region (8), at least one inlet (12) for introducing a fuel into the mixing region (8) and at least one inlet (10, 13) for introducing steam into the mixing region.

Abstract: A gas turbine combustor, which can enable only a swirler and a heat shield to be readily taken out from the combustor upon the exchange the swirler and the heat shield. The gas turbine combustor includes: a combustion cylinder; a fuel injection unit for supplying fuel to a head portion of the combustion cylinder; a support member configured for allowing the fuel injection unit to be supported by the combustion cylinder; and a heat shield adapted for heat-insulating the support member from combustion gas in the combustion chamber, wherein the fuel injection unit includes a fuel injection valve for injecting the fuel, and a swirler for supplying compressed air to the fuel injected from the fuel injection valve while swirling the compressed air. The swirler and heat shield are connected together, constituting a swirler unit. This swirler unit is detachably attached to the support member via a fastening member.

Abstract: Systems and methods provide for wear reduction in a combustion system of a gas turbine. A system for wear reduction includes: at least one substantially H-shaped block, the substantially H-shaped block being configured to secure a transition piece of a gas turbine combustor to a support piece; a first insert including a tungsten carbide in a metal matrix, the metal matrix being selected from a group including cobalt and nickel; and a brazing material which is used in brazing the first insert to the at least one substantially H-shaped block in at least one location on an interior wear surface of the at least one substantially H-shaped block.

Abstract: A method and system suitable facilitating the connection and disconnection of a thrust reverser assembly to a fan case of a nacelle of a gas turbine engine. The method and system entail operating a clamping system to simultaneously engage and disengage flanges associated with the fan case and flanges associated with the fixed structure of the thrust reverser assembly. The clamping system includes a plurality of over-center clamping mechanisms, each of which is movable to simultaneously clamp together complementary flanges of the thrust reverser assembly and the fan case.