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: A mounting arrangement for an aircraft engine includes a mounting structure attached to an aircraft body and not directly connected to an engine core of the aircraft engine, a propulsor mounted to the mounting structure, and a thrust reverser mounted directly to either the propulsor or mounting structure.

Abstract: A gas turbine engine includes a first annular portion that is stationary and adapted for partially surrounding an engine core. The first annular portion includes a fore pylon connecting portion. The gas turbine engine also includes a rail coupled to the fore pylon and extending in the aft direction from the first annular portion. The gas turbine engine also includes a second annular portion, arranged aft of the first portion and coupled to the rail. The second annular portion is movable along an engine core centerline between a closed position and at least one open position. The second annular portion is configured to engage the first annular portion in the closed position, thereby providing access to the engine core. The gas turbine engine further comprises a thrust reverser arranged in the second annular portion.

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: The combustion device (1) of a gas turbine includes burners (2) connected to a front plate (5) of a combustion chamber (3). The front plate (5) has, spaced apart from one another, a front sheet (8) and an impingement sheet (7) with aligned holes (11, 12) housing the burners (2). A piston ring (15) is provided between the front sheet (8) and impingement sheet (7) to seal the holes (11, 12). The axial length of the border of the hole (11, 12) of the front sheet (8) and/or impingement sheet (7) is longer than the thickness of the corresponding front sheet (8) and/or impingement sheet (7).

Abstract: Transition duct assemblies for turbine systems and turbomachines are provided. In one embodiment, a transition duct assembly includes a plurality of transition ducts disposed in a generally annular array and comprising a first transition duct and a second transition duct. Each of the plurality of transition ducts includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct assembly further includes an aerodynamic structure defined by the passages of the first transition duct and the second transition duct. The aerodynamic structure includes a pressure side, a suction side, and a trailing edge, the trailing edge having a modified aerodynamic contour.

Abstract: A gas turbine engine having a fan and a bypass duct for discharging airflow from the fan to generate engine thrust. The bypass duct includes a first component and a second component each having a gas washed surface, the first component being positioned adjacent the second component with a gap therebetween. A sealing strip is positioned in the gap between the first component and the second component distal to the gas washed surfaces of the first and second components. A sealant is positioned between the first and second component proximal to the gas washed surfaces of the first and second components, so as to seal the gap between the first component and the second component from airflow through the bypass duct.

Abstract: A gas turbine enclosure apparatus is provided and includes a frame defining an interior, gas turbine equipment disposed in the interior, the gas turbine equipment being pre-assembled at a first site and connectable with a gas turbine disposed at a second site and wall panels connectable with the frame. The apparatus being disposable proximate to the gas turbine such that the gas turbine equipment is exposed to the gas turbine and a portion of the wall panels form a portion of an outer enclosure of the gas turbine and the gas turbine equipment.

Abstract: A gas turbine engine includes a combustor, a first turbine section in fluid communication with the combustor, a second turbine section in fluid communication with the first turbine section, and a mid-turbine frame located axially between the first turbine section and the second turbine section. An oil system includes a first oil system component that houses oil. A first connector mechanically connects the first oil system component to the mid-turbine frame.

Abstract: An aircraft nacelle including a pipe (32), a peripheral wall (34), a lip (36), and a front frame (38) connecting the peripheral wall (34) and the pipe (32) and forming with the lip (36) a space (50) in which the hot air that is provided for frost treatment can circulate, with the pipe (32) including a coating (44) for the composite material acoustic treatment, is characterized in that the nacelle includes at least one element (52) made of a heat-conducting material inserted between the lip (36) and the pipe (32) ensuring the continuity of the aerodynamic surfaces of the lip (36) and the pipe (32) and the propagation of heat from the space (50) toward the rear of the nacelle, whereby the at least one element (52) includes a coating (54) for acoustic treatment that is made of a heat-resistant material.

Abstract: A transition piece aft frame assembly is provided, and includes a transition piece aft frame and a heat shield. The transition piece aft frame has an aft face. At least a portion of the aft face is exposed to an exhaust gas stream. The heat shield is connected to the transition piece aft frame. The heat shield is oriented to generally deflect the exhaust gas stream away from the aft face of the transition piece aft frame.

Abstract: A system including an air intake subsystem having a housing is provided. A coalescer having a frame is disposed inside the housing. The system includes a mounting bracket, and a release mechanism coupled to the frame and the mounting bracket. The release mechanism is adapted to selectively release the frame.

Abstract: Devices and methods for supporting an accessory (32) in a gas turbine engine (10) are disclosed. The accessory (32) may have an interface (36) for coupling to a shaft (24) of the gas turbine engine (10). The device may comprise a first support (38) configured to support a first portion (34) of the accessory (32) proximal the interface (36) and a second support (40) configured to support a second portion (35) of the accessory (32) distal from the interface (36). The second support (40) may be configured to provide a lower resistance to relative displacement between the accessory (32) and structure (42) of the engine (10) than does the first support (34).

Abstract: A supporting structure for a gas turbine engine includes at least one annular member, and a plurality of circumferentially spaced elements that extend in a radial direction of the annular member and that are connected to the annular member. At least one of the circumferentially spaced elements has an airfoil shape in cross section. The annular member includes at least two substantially flat panel segments that are arranged side by side in a circumferential direction of the annular member. The panel segments are connected to each other in a connection region that extends along facing end parts of the two panel segments. The airfoil shaped element is connected to the annular member at the connection region. A mean camber line of the airfoil shaped element at least along a portion of its length is inclined in relation to an axial direction of the annular member, and the connection region at least along a portion of its length is inclined in relation to the axial direction of the annular member.

Abstract: An aircraft nacelle includes at its outside wall a cowl moveable relative to the nacelle so as to block or unblock an opening. The cowl includes an articulation and locking/unlocking elements distant from the upstream edge of the cowl that include elements for limiting the appearance of scooping phenomena. The limiting elements include at least one lock integral with the rest of the nacelle or respectively the cowl that can occupy a locked state in which part of the lock interferes with the cowl or respectively the rest of the nacelle and prevents a deformation that includes a radial component of the cowl, and another free state in which the part no longer interferes with the cowl or respectively the rest of the nacelle and allows the opening movement of the cowl, with the switch from one state to the next being generated by a gas pressure variation.

Abstract: A gas turbine engine high pressure rotor with first and second high pressure turbine stages include first and second stage disks having first and second stage disk bores and a single tie rod therethrough. First and second bore annular flowpaths are radially located between first and second stage disk bores and tie rod. A means for increased cooling and/or heating in second stage disk bore is axially located within second stage disk bore. The means may include an airflow accelerator such as one or more annular ribs on the tie rod. A bore annular cross-sectional flow area between the second stage disk hub and the ribs may be substantially smaller than between the second stage disk hub and the tie rod. An axially unobstructed inlet into the second bore annular flowpath allows fully axially flowing and axially unobstructed flowing of second stage bore cooling air into inlet.

Abstract: A system includes a combustor transition piece configured to mount between a combustor and a turbine of a gas turbine engine. The combustor transition piece includes a hollow body with an internal flow passage extending from an upstream end portion to a downstream end portion. The combustor transition piece also includes an aft frame coupled to the downstream end portion of the hollow body, an aft flange coupled to the aft frame, and a mounting lug removably coupled to the aft frame at a first joint. The mounting lug is configured to removably couple to a mounting bracket at a second joint.

Abstract: An exhaust includes a wall that has a first composite material having a first coefficient of thermal expansion and a second composite material having a second coefficient of the thermal expansion that is less than the first coefficient of thermal expansion.

Abstract: A device (1) for connecting a fixed turbine engine portion (2) and a distributor (3) foot (31) of a turbine engine turbine, comprising: a ring-shaped body (11) adapted so as to be connected and secured to the fixed turbine engine portion (2), and two ring-shaped upstream (12) and downstream (13) brackets secured to the body (11), both brackets (12, 13) being adapted so as to clamp the foot (31) of the distributor, characterized in that at least one bracket (12, 13) is interrupted by forming an expansion or retraction slot (121, 131), the slot (121, 131) allowing an expansion of the bracket (12, 13) tending to close the slot (121, 131) or retraction of the bracket (12, 13) tending to further open the slot (121, 131), under the effect of a difference in internal temperatures of the device.

Abstract: A fuel nozzle assembly for use in a combustor apparatus of a gas turbine engine. An outer housing of the fuel nozzle assembly includes an inner volume and provides a direct structural connection between a duct structure and a fuel manifold. The duct structure defines a flow passage for combustion gases flowing within the combustor apparatus. The fuel manifold defines a fuel supply channel therein in fluid communication with a source of fuel. A fuel injector of the fuel nozzle assembly is provided in the inner volume of the outer housing and defines a fuel passage therein. The fuel passage is in fluid communication with the fuel supply channel of the fuel manifold for distributing the fuel from the fuel supply channel into the flow passage of the duct structure.

Abstract: The invention relates to a carrier ring for a high-pressure gas turbine of a gas turbine plant, which has a high-pressure combustion chamber upstream of the high-pressure gas turbine, a compressor upstream of the high-pressure combustion chamber, a low-pressure combustion chamber downstream of the high-pressure gas turbine, a low-pressure gas turbine downstream of the low-pressure combustion chamber, and a rotor that carries rotor blades for the compressor, for the high-pressure gas turbine, and for the low-pressure gas turbine. The carrier ring carries guide blades and/or heat shields of the high-pressure gas turbine and can be fastened to the high-pressure combustion chamber. The installation of the carrier ring can be simplified by segmenting the carrier ring at least in the area of the guide blades thereof and/or of the heat shields thereof in the circumferential direction, wherein the segmented carrier ring has at least two ring segments that carry the guide blades and/or the heat shields.

Abstract: The present application provides a fuel plenum for a micro-mixer combustor. The fuel plenum may include a first assembly plate with a first assembly plate aperture, a second assembly plate with a second assembly plate aperture, a fuel tube extending through the first assembly plate aperture of the first assembly plate and the second assembly plate aperture of the second assembly plate, and an installation insert positioned between the fuel tube and the first assembly plate aperture.

Abstract: A suspension assembly for an aircraft turbojet engine includes a pylon and connecting rods for taking up the thrust of the turbojet engine which is connected to the pylon. In particular, the suspension assembly includes an insulator which is separate from the pylon and thermally insulates the connecting rods from the turbojet engine.

Abstract: A method and apparatus are disclosed for a gas turbine spool design combining metallic and ceramic components in a way that controls clearances between critical components over a range of engine operating temperatures and pressures. In a first embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud and separated by a small clearance gap. The ceramic rotor is connected to a metallic volute. In order to accommodate the differential rates of thermal expansion between the ceramic rotor and metallic volute, an active clearance control system is used to maintain the desired axial clearance between ceramic rotor and the ceramic shroud over the range of engine operating temperatures. In a second embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud which is part of a single piece ceramic volute/shroud assembly.

Abstract: An assembly for mounting a turbine engine to a pylon includes a mounting beam, a plurality of fasteners, a first mounting linkage and a second mounting linkage. The mounting beam includes a mount beam fitting that extends axially between a first mount beam end and a second mount beam end, and a mount beam flange that extends radially out from the mount beam fitting at the first mount beam end. A first fastener aperture extends radially through the mount beam fitting, and a second fastener aperture extends axially through the mount beam flange. The fasteners connect the mounting beam to the pylon. A first of the fasteners is mated with the first fastener aperture, and a second of the fasteners is mated with the second fastener aperture. The first mounting linkage connects the first mount beam end to a first engine attachment. The second mounting linkage connects the second mount beam end to second engine attachments.

Abstract: A combustor heat shield assembly comprises a circumferential array of heat shield panels individually mounted to an inner surface of a combustor shell. Each heat shield panel has a sealing rail extending from a back side thereof and a plurality of bolted connections securely holding the heat shield panel on the combustor shell with the sealing rail in sealing contact with the inner surface of the combustor shell. Each pair of adjacent heat shield panels comprises first and second panels having adjoining lateral edges, the first panel having a boltless area on its back side at a location adjacent to its adjoining lateral edge. The second panel has a first one of its bolted connections provided adjacent to its adjoining lateral edge and in facing relationship with the boltless area of the first panel.

Abstract: A gas turbine engine assembly includes first and second annular members having different first and second thermal expansion coefficients connected together with dual arm V brackets. Brackets include first and second arms angularly spaced apart from a bracket centerline and extending axially away from bracket bases attached to a first one of the first and second annular members. Arms are attached to a second one of the first and second annular members. A turbine frame includes struts extending between outer and inner rings. An annular mixer and centerbody substantially made from a ceramic matrix composite materials is connected to and supported by the outer and inner rings with first and second sets respectively of the dual arm V brackets. Bracket bases of the first and second sets are attached to the outer and inner rings respectively. Arms of the first and second sets are attached to mixer and centerbody respectively.

Abstract: A gas turbine engine has a rear mounting structure incorporating a mounting apparatus attached to a bypass duct wall with a link device of six rods for transferring core portion related inertia-induced loads, from an inner case of the core portion in a short circuit across an annular bypass air passage to the bypass duct wall.

Abstract: The mounting system can be used in mounting a turbofan gas turbine engine having a nacelle including two halves, each half having a hinged end hingedly connected to the pylon and a free end lockingly engageable with the engine structure. The mounting system can include a primary connection connecting the engine structure to the pylon, and a flexible connection provided between the hinged connections and the pylon, the flexible connection being elastically deformed when a significant force is exerted in a transversal plane upon one of the halves in the closed position.

Abstract: A loading assembly for a turbine system is disclosed. The loading assembly includes a transition duct and a load member. The transition duct extends between a fuel nozzle and a turbine section, and has an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The load member extends from the transition duct and is configured to transfer a load between the transition duct and an adjacent transition duct along at least one of the longitudinal axis, the radial axis, or the tangential axis.

Abstract: A gas turbine engine includes a tangential on-board injector (TOBI) fluidly connected to a compressor section. A diffuser case structurally supports a combustor section and the tangential on-board injector via at least one low thermal mass joint.

Abstract: An interface (10) is provided between a combustor basket (12) and a transition (14) of a gas turbine engine (16). The interface (10) includes a tongue (18) formed in one of the combustor basket (12) and the transition (14), and a groove (20) formed in another of the combustor basket (12) and the transition (14). The tongue (18) is received within the groove (20) to define the interface (10). A sealing arrangement is also provided and includes an annular combustion outlet (12??) and an annular transition inlet (14??), with a labyrinth seal (72??) positioned between an end (68??) of the annular combustion outlet and an end (70??) of the annular transition inlet to form a labyrinth path (34??) there between. An inner diameter (24??) of the annular combustion outlet (12??) is axially aligned with an inner diameter (22??) of the annular transition outlet (14??).

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: A coupling support member including a pair of divided pieces is placed in a coupling part between a vane proximal end portion of a guide vane and an attachment flange, and the pair of divided pieces are joined to the vane proximal end portion from both the sides in the vane thickness direction. A linear protrusion is formed on one of joint surfaces of the vane proximal end portion to the pair of divided pieces. A groove engaged with the linear protrusion formed in the vane proximal end portion is formed on one of respective joint surfaces of the pair of divided pieces of the coupling support member. The vane proximal end portion is held between the pair of divided pieces of the coupling support member, by the fastening force that is applied to the pair of divided pieces of the coupling support member from both the sides in the vane thickness direction. It is possible to obtain a high structural strength while contributing to a reduction in weight of a jet engine.

Abstract: A turbofan gas turbine engine is provided having a unique power off-take shaft and gear system. Other gas turbine engine types are also contemplated herein. Two power off-takes are provided, one each for the low pressure spool and high pressure spool. The power off-takes extend across a core flow path of the turbofan engine between the low and high pressure shafts to a fan frame of the turbofan. A drive gear is provided near the front end of the high pressure shaft, and another drive gear is provided on the low pressure shaft near the drive gear for the high pressure shaft. Both gears are located in a sump of the gas turbine engine. The power off-take shafts are coupled to the drive gears. Two power devices are coupled to the power off-take shafts and are located in the fan frame. The power devices can be electric generators or motors.

Abstract: In a gas turbine engine, a fan rotor and a compressor rotor are connected to a joint which is attached to a shaft. The fan rotor is connected to the joint by a plurality of fasteners extending through mounting openings in the joint and through apertures in the fan rotor. A standard retaining ring is attached to the joint to cover a portion of an enlarged head of each of the respective fasteners in order to prevent removal of the respective fasteners from the mounting openings in the joint before the respective fasteners are connected to the fan rotor.

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 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 burner arrangement is provided. The burner arrangement includes a support and at least two fuel nozzles attached to the support in the direction of flow, with each fuel nozzle including a support-side section which includes a contact surface on the support side with which it rests on a supporting surface of the support, with at least two fuel nozzle tips embodied in one piece extending out from the support-side section in the direction of flow and the support-side contact surface including at least two extension parts projecting in the direction of the support, with the extension parts each embodying a channel through which fuel is fed in each case to the fuel nozzle tips through passages which are arranged in the support-side contact surface of the support-side section.

Abstract: An aft aerodynamic fairing for a pylon for attaching an aircraft propelling assembly, including a caisson structure, a heat-shield floor, as well as a support structure for this floor comprising linking parts arranged to link two side edges of this floor to the caisson structure. The linking parts of each side edge of the floor are configured so as to allow travelling motion of the side edge relative to the caisson structure in a corresponding direction included in a plane orthogonal to a longitudinal direction of the fairing. The linking parts form a sliding link between the floor and the caisson structure, reducing the stresses induced by the deformations of these elements in use.

Abstract: A system for supporting bundled tube segments within a combustor includes an annular sleeve that extends circumferentially and axially within the combustor, a support lug that extends radially inward from the annular sleeve and an annular support frame that is disposed within the annular sleeve. The annular support frame includes an inner ring portion, an outer ring portion and a plurality of spokes that extend radially between the inner and outer ring portions. The inner ring portion, the outer ring portion and the plurality of spokes define an annular array of openings for receiving a respective bundled tube segment. The inner ring portion is connected to each bundled tube segment and the outer ring portion is coupled to the support lug.

Abstract: A compressor blade mounting arrangement includes a compressor blade having an airfoil, a base attachment and a platform disposed between the airfoil and the base attachment. The base attachment includes a first root portion extending radially inwardly from the platform and a second root portion extending radially inwardly from the platform, the first root portion and the second root portion defining a hollow portion therebetween. The base attachment also includes a compression plate disposed within the hollow portion and in fitted engagement with the first root portion and the second root portion. The base portion further includes at least one tab protruding from at least one of the first root portion and the second root portion. The compressor blade mounting arrangement also includes a wheel including a channel extending from a base wall to a rim of the wheel, the channel having a varying opening width.

Abstract: A mounting arrangement for a gas turbine engine includes a pylon mount structure that has a fore mount portion and an aft mount portion. A first engine mount beam is integrally formed with the fore mount portion as a single-piece component, and a second engine mount beam is integrally formed with the aft mount portion as a single-piece component. The first engine mount beam is attachable to the engine case at a fore attachment interface, and the second engine mount beam is attached to the engine case at an aft attachment interface.

Abstract: The invention concerns a gas turbine engine component (27) comprising an outer ring (21), an inner ring (20), a plurality of circumferentially spaced elements (22) extending between the inner ring (20) and the outer ring (21), wherein an annular load transfer structure (23) extends circumferentially along an inner side of the inner ring (20) and also inwards in a radial direction of the component (27). The first portion (23a), at least along a part of the circumference, is inclined in the radial direction in relation to a second portion (23b), wherein the two inclined portions (23a, 23b) are connected in a connection zone (26) between the inclined portions (23a, 23b), and wherein a radial and/or axial position of the connection zone (26) varies along the circumference such that the radial/axial position of the connection zone in a location that circumferentially corresponds to a first element (22a), i.e.

Abstract: The present invention relates to a device (20) for installing and/or removing a pre-cooler (16) of an aircraft engine (10). Said device comprises, in a position of use, a mounting flange (22) supported in use by a pylon of an aircraft, at least a floor (68) supporting a pair of guiding arms (24), a ramp (26), a basket (32) for receiving said cooler (16) and adjustable support means (30) of said cooler for removing or mounting said cooler on pylon (16).

Abstract: The invention relates to a rear section (11) of an aircraft nacelle, that comprises two halves (13a, 13b) defining: a central portion (C) for receiving a turbojet engine (7), a cool-air annular passage (31) provided around said central portion (C), and at least one six-hour cavity (15) provided under said central portion (C). The rear section is characterized in that it comprises at least one duct (29a, 29b) for the fluidic communication between said annular passage (31) and said six-hour cavity (15) for maintaining the temperature inside the six-hour cavity (15) within a relatively low range.

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 combustor for a gas turbine engine is disclosed. The combustor is described as comprising a dome plate coupled to a liner thereof, with at least one heat shield comprised of a ceramic matrix composite coupled at the aft end of the dome plate. Also described is a method for assembling a combustor for a gas turbine engine, including releasing a metal alloy heat shield from a dome plate and providing a ceramic matrix composite heat shield as replacement.

Abstract: An intermediate support frame (68) that spans an inner diameter of a support ring (48) of a gas turbine combustor cap assembly (24) at a position intermediate the length of the support ring. The intermediate support frame may have a central encirclement (72) that receives a central fuel pre-mix tube (44) of the combustor cap assembly, and may further have a circular array of outer stabilization rings (70) that each receive a respective outer pre-mix tube (42). The central pre-mix tube may be affixed to the central encirclement (72), for example by welding. The outer pre-mix tubes may be slidably engaged in the outer stabilization rings (70), providing lateral stability while allowing differential thermal expansion. The intermediate support frame may have holes (74) for coolant passage, and perimeter tabs (76) for attachment to the support ring (48).

Abstract: A gas turbine combustion chamber has a starter film for cooling the combustion chamber wall, and a combustion chamber head, into which cooling air can be introduced and which is confined to the combustion chamber by a heat shield (5). A base plate (2) is arranged at a certain distance from the heat shield (5) and the base plate (2) is provided with several openings (6) in its rim area for passing the cooling air. Center axes (17) of the openings (6) are inclined at a shallow angle (?) relative to the combustion chamber wall (4).