Abstract: A turbo machine is provided that includes a turbine nozzle having a wall. The wall provides spaced apart first and second surfaces. A combustor includes a liner having an annular lip that engages one of the first and second surfaces. A floating ring seal is supported by the liner in a slip-fit relationship, for example, using a retainer. The floating ring seal engages the other of the first and second surfaces. The floating ring seal is slidably moveable relative to the liner in a floating direction in response to movement of the liner in a radial direction.

Abstract: One form of the present invention contemplates an apparatus including a gas turbine engine core, wherein a gearbox accessory is mounted to the engine core and is substantially enclosed by an inner fairing. The apparatus can further include a nacelle having a plurality of panels, wherein the nacelle panels substantially enclose and overlap at least a portion of the inner fairing. The apparatus can further include a bypass duct defined between an exterior surface of the inner fairing and an interior surface of at least some of the nacelle panels. In additional and/or alternative forms a number of frame rods can also be used internal to the nacelle or nacelle panel. In some forms the frame rods can be connected between a front and rear frame.

Abstract: The present invention provides a hula seal for use with a combustor. The hula seal includes a number of legs that define a number of slots. The slots may include a number of expansion slots.

Abstract: A quick disengaging field joint connects a first component of an exhaust system of a gas turbine engine to a second component of the exhaust system. The field joint includes a pair of opposed stepped liners connected via exterior-facing connecting flanges. The field joints can be disassembled entirely from outside the exhaust housing without requiring access to the interior of the exhaust housing.

Abstract: A transition duct having a thermally free aft frame and being capable of adjusting the natural frequency is disclosed. The aft frame is capable of permitting movement due to thermal gradients with the transition duct. The transition duct utilizes a spring plate located adjacent to an aft mounting bracket, where the spring plate, based on its thickness can either increase or decrease a frequency of the transition duct. Such an arrangement ensures that the transition duct natural frequency does not coincide with or cross other critical engine and/or combustor frequencies.

Abstract: An exemplary turbine engine link assembly includes a link extending longitudinally from a rod end and terminating at a hemispherical end. The rod end is secured to an engine liner or an engine casing. The hemispherical end is biased toward a corresponding hemispherical recess in the engine liner or in the other of the engine liner or the engine casing.

Abstract: An air cycle machine includes two turbines and a compressor mounted on an integral shaft. The integral shaft includes a plurality of shaft sections that are welded together and machined in a single set-up process into a desired shaft shape to provide highly aligned bearing surfaces. The shaft includes three stops that cooperate with three fasteners to secure the two turbines and the compressor on the shaft.

Abstract: A lower cost combustor is formed from a series of sheets of metal that are secured by mechanical fasteners. An increase in diameter of the combustor is achieved through the provision of securing rings to which the sheets of metal are secured.

Abstract: A duct having a continuous C-shaped cross section slidably attached to a strut supporting an aircraft engine. The duct may comprise an inner wall, an outer wall, a first end wall, a second end wall, at least one slider, a thrust reverser, and an outer cowl panel of a thrust reverser. The inner wall may comprise a forward section and an aft section. The aft section may be integral with the first end wall, the second end wall, the outer wall, and the at least one slider. The forward section may comprise a right half and a left half which may be pivotally attached to the strut and may pivot to an open position when the rest of the fan duct is slid aftward away from the engine. The duct, excluding the forward section of the inner wall, may slide aftward on its at least one slider, providing engine access.

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

Abstract: A boss and orifice plate assembly comprising an annular boss adapted to be secured in a hole formed in a combustor component, said boss formed with an annular seat supporting a replaceable orifice plate, and an annular retaining ring groove adjacent said seat, said seat extending radially inwardly of said annular retaining ring groove; and a retaining ring seated in said retaining ring groove and at least partially and resiliently engaged between a surface of said groove and a surface of said orifice plate.

Abstract: A combustor liner for a gas turbine engine including first and second annular liner portions engaged to one another with an interference fit. The first liner portion includes at least one anti-rotation feature extending therefrom and engaging the second liner portion and having a threaded hole defined therethrough at least substantially perpendicularly to an adjacent radially extending surface of the second liner portion for threadingly receiving a pushing tool for pushing against that surface.

Abstract: A vane fixing apparatus for removably fixing a vane plate (14) to a web (12) is described. The vane fixing apparatus comprises a mounting bracket (16) attached to the vane plate (14), and a mounting block (20,30) held on the web (12) by at least one fastener (18), wherein the mounting block (20,30) is moveable between a first configuration in which the mounting bracket (16) may engage and disengage the mounting block (20,30) and a second configuration in which the mounting block (20,30) securely retains the mounting bracket (16) relative to the web (12), and wherein the mounting block (20,30) is moveable between the first and second configurations by operation of the fastener (18).

Abstract: A flame holder for an afterburner duct of a jet engine, including an arm in the form of a gutter forming a cavity, a shield for protecting the cavity of the arm from heat, and an air supply baffle housed in the cavity, is disclosed. The arm, the protective shield, and the air supply baffle are held together by a one-piece spacer shoe.

Abstract: An access conduit for access to the interior of a combustor for the washing of the turbine components of a gas turbine engine. The access conduit is secured between a gas generator case and a combustion chamber liner of the combustor. The access conduit has an open outer end accessible from an outer face of the gas generator case and an open inner end exiting in a combustion chamber through the liner. The access conduit has one or more openings disposed in a combustor gap between the gas generator case and the liner of the combustor. A plug is removably secured in the open outer end and extends into the access conduit to obstruct the one or more openings.

Abstract: Disclosed is an apparatus configured to seal with a turbine blade stage of a gas turbine. The apparatus includes an outer shroud coupled to an inner shroud and configured to circumferentially surround the turbine blade stage. The inner shroud is configured to circumferentially surround the turbine blade stage to seal with the turbine blade stage and includes an attachment element configured to be inserted into the outer shroud to couple the inner shroud to the outer shroud.

Abstract: A securing system is provided for securing a component to a gas turbine engine casing. The system has a core engine casing formed from a front casing portion and a rear casing portion. The casing portions have respective casing flanges at which the casing portions are detachably joined together. The system further has a component flange extending from a component. The component flange is configured so that, when the casing portions are joined together, the component flange is trapped between the casing flanges. One of the casing flanges projects inwards from the casing to form an airtight seal with the component flange on the inside of the casing. The other of the casing flanges projects outwards from the casing to form an airtight seal with the component flange on the outside of the casing.

Abstract: Twin-flow turbo-jet engine, comprising a cylindrical casing externally defining an annular flow space for a secondary flow, wherein the cylindrical casing is formed from a lattice framework and from removable cowling panels fixed to the framework, said framework comprising an upstream annular flange for fixing to an intermediate housing, a downstream annular flange for connecting to an exhaust housing and rigid beams parallel or inclined relative to the axis of the turbo-jet engine, connecting the two flanges to one another.

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: 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 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: An apparatus for internally mounting a fuel manifold within a surrounding casing in a gas turbine engine which includes a fuel manifold and support pins which are displaceable relative to mounting portions on the fuel manifold between a retracted position, wherein the fuel manifold is able to be inserted within a casing of the engine, and an extended position, wherein the support pins project outwardly from the fuel manifold for engagement with the casing.

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 gas turbine engine has a casing component, such as a mid turbine frame system, which includes an outer case surrounding a spoke or strut casing. Load transfer spokes of the spoke casing are detachably connected to the outer case and radially extend between the outer case and an inner case of the spoke casing. A bearing housing is supported by the inner case of the spoke casing. An arrangement for radially centring the bearing within the outer case is provided, the arrangement including at least three radial locators mounted to the outer case.

Abstract: A combustor assembly having a transition piece and at least one orifice assembly in the transition piece, the orifice assembly comprising: a boss having an outside periphery and an inside periphery, the inside periphery including an annular seat and an upstanding flange formed with an annular, inwardly facing retaining ring groove, the boss fixed within an opening in the transition piece; an orifice plate having a bottom surface that is adapted to be received on the annular seat; and a retaining ring located in the retaining ring groove and at least partially engaged with the orifice plate.

Abstract: An engine cowl door assembly is provided for fuselage mounted turbine engines to minimize manufacturing costs and part counts. In one embodiment, the engine cowl door assembly includes at least one non-handed engine cowl door, a pylon apron coupled to the fuselage of an aircraft and providing an attachment point for a first portion of the engine cowl door and a fairing coupled to the turbine engine and providing a latching point for a second portion of the engine cowl door to the turbine engine. The pylon apron, positioned on an inboard side of the turbine engine, and the fairing, positioned on an outboard side of the turbine engine, are identically shaped to provide interchangeability of the engine cowl door from a turbine engine mounted on a first side of an aircraft fuselage to a turbine engine mounted on a opposed second side of an aircraft fuselage.

Abstract: A power system includes a gas turbine engine. The gas turbine engine may include a first portion, a second portion, and a longitudinal axis. The power system may also include a first moveable support engaged to the first portion of the gas turbine engine. The first moveable support may be operable to at least partially constrain movement of the first portion of the gas turbine engine during operation of the gas turbine engine. Additionally, the first moveable support may be operable to at least partially support the first portion of the gas turbine engine as the first portion of the gas turbine engine is moved away from the longitudinal axis independently of the second portion of the gas turbine engine.

Abstract: An annular turbomachine combustion chamber comprising an inner wall, an outer wall, and a chamber end wall disposed between said inner and outer walls in the upstream region of said chamber, the chamber end wall presenting a central portion and an outer assembly flange, and the outer wall presenting an upstream assembly flange, the chamber end wall and the outer wall being assembled together by their assembly flanges, in which the free edge of said outer assembly flange is curved inwards. In addition, the outer face of the outer assembly flange presents a shoulder.

Abstract: A method of constructing, assembling and disassembling a mandrel for fabricating a monolithic composite nacelle panel for use on an airplane comprises attaching a first fairing bar including an exposed ball roller assembly to a tilting plate; attaching a first outer section to said first fairing bar; attaching a second outer section to said first fairing bar; attaching a center section to said first outer section and said second outer section; attaching a nozzle section to said first fairing bar and positioned adjacent an inner wall of said first outer section, an inner wall of said center section, and an inner wall of said second outer section; attaching a second fairing bar to said first outer section, said center section, and said second outer section; and attaching a cavity section to said second fairing bar.

Abstract: The invention relates to a jet engine nacelle intended to equip an aircraft, formed from at least first and second half-shells and equipped with at least one locking system designed to lock one of the half-shells with respect to the other, the locking system being actuable between a locked state and an unlocked state by means of an actuating device (14). The actuating device (14) comprises a fixed body (29) in which an actuating rod (16) is translatably and rotatably mounted with respect to its longitudinal axis, the rod (16) having a first end equipped with a handle (17) and a second end connected to the locking system, the fixed body (29) having a track (30) cooperating with a follower element (24) projecting from the rod (16), the track (30) having a central portion (31) extending parallel to the axis (A) of the rod (16), on either side of which are respectively formed first and second end portions (32, 33) extending substantially perpendicularly to the axis of the rod (16).

Abstract: An arrangement for sealing an opening formed in a wall for receiving a member. The arrangement comprises a sealing member with a passage, and the sealing member is moveably mounted on the wall adjacent to the opening such that the sealing member passage is substantially co-axial with the opening in the wall. Biasing means for biasing the sealing member towards the wall are also provided.

Abstract: A gas turbine engine has a compressor assembly and a turbine assembly rotationally mounted on a shaft, the turbine assembly being driven by hot gases discharged from a combustion chamber disposed between the compressor and turbine assemblies, the compressor having a centrifugal impeller for pressurizing and impelling air into the combustion chamber. The engine also includes an impeller shroud covering the bladed portion of the centrifugal impeller, the impeller shroud having a support bracket having a thin and curved load-isolating profile for supporting a strut that secures the impeller shroud to a case of the engine.

Abstract: A fuel nozzle for a gas turbine engine comprises a nozzle body defining a passage extending therethrough and a flange around the nozzle body extending radially and outwardly from the nozzle body. The flange defines a mounting face thereof and a plurality of mounting holes in the mounting face extending through the flange. The mounting face includes a contacting land protruding from the mounting face for abutting a mounting surface of a support structure of the engine when the fuel nozzle is mounted to the support structure.

Abstract: A method of assembling a combustor for use in a turbine engine is disclosed. The method includes providing a dome assembly ring including a plurality of assembly ring openings, positioning a plurality of elongated rings on the dome assembly ring, providing a cowl assembly including an inner cowl portion and an outer cowl portion that are fabricated from sheet metal, and coupling the inner and outer cowl portions to the dome assembly ring such that each of the plurality of elongated rings is coupled to the dome assembly ring.

Abstract: A one-piece combustion chamber having an end wall subdivided into sectors. The combustion chamber comprises a chamber end wall interconnecting the inner and outer walls, and it includes windows closed by removable walls, each carrying a plurality of injector systems.

Abstract: A compound bracket system (400) for an internal combustion engine (100) includes a first bracket (112) that is rigidly connected to a crankcase (102) of the internal combustion engine (100). The first bracket (112) forms at least one mounting pad (310) that is arranged to connect to and support a first engine component (114). A second bracket (120) is connected to the crankcase (102) through at least one additional engine component (110). The second bracket (120) forms a component cavity (208) that is arranged to accept and support a second engine component (118). The first bracket (112) forms an interconnection pad (314), and the second component forms a strut (216) having a mounting tab (218). The mounting tab (218) is advantageously connected to the interconnection tab (314) with a fastener (422), thus increasing the rigidity of the second bracket (120) as mounted on the engine (100).

Abstract: A shroud for a combustion chamber bottom designed to cover fuel injectors is provided with drillings on at least one of its sides to open up the cavity within the shroud and reduce noise that it produces and combustion instabilities. The drillings also have the effect of reducing instabilities and non-uniformity of the airflow around the shroud.

Abstract: A device for supporting and housing auxiliaries in a bypass turbojet is disclosed. The device includes two coaxial case rings arranged one inside the other and connected by tubular radial arms through which fluid ducts and electric cables pass. At least one of the radial arms includes on a lateral face a detachable panel whose removal allows access to turbojet equipment situated radially inside the internal case ring in alignment with the radial arm.

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

Abstract: The combustion chamber includes a chamber endwall perforated by at least one passage hole. An injection system is mounted so that it can slide diametrically with respect to the passage hole. A deflector is mounted on the chamber endwall. The deflector includes a fitting ring which includes first tenons. A sleeve which is coaxial to the fitting ring includes second tenons. The first and the second tenons allow the passage of the first tenons between the second tenons and then the engagement of the first tenons behind the second tenons. The injection system is slidably mounted in the sleeve. Antirotation means are provided to prevent a rotation of the injection system with respect to the sleeve.

Abstract: A method for assembling a combustor for use in a turbine engine is provided. The method includes providing at least one support leg, forming at least one opening in the support leg, wherein the at least one opening has a perimeter defined only by a plurality of non-linear segments, and coupling the support leg to the combustor.

Abstract: A sacrificial liner is provided for an injector access aperture in a combustor casing. The liner has an outer annular surface and an eccentric inner annular surface. The liner serves to protect the combustor outer casing from wear by a series of piston rings mounted in a seal carrier. The eccentricity of the surfaces prevent excess contact pressure between the piston rings and the seal carrier to improve the life of these components.

Abstract: An exhaust nozzle for a gas turbine engine according to an exemplary aspect of the present invention includes: an outer cowl; a cowl interface ring attached to the outer cowl; an acoustic nozzle; an acoustic nozzle interface ring attached to the acoustic nozzle, the acoustic nozzle interface ring attachable to the cowl interface ring; and a removable nozzle section attachable to the acoustic nozzle interface ring and the cowl interface ring.

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 load transfer unit, a torque box rotatably positioned within the load transfer unit, and a plurality of struts. The load transfer unit has a first locking element and combines the first load and the second load into a combined load. The torque box has a second locking element that is engagable with the first locking element of the load transfer unit. The plurality of struts are connected between the torque box and the mount, and transfer the combined load from the torque box to the mount. The first locking element and the second locking element are at least one of a rib or a groove.

Abstract: A method of dismantling a turbomachine includes opening a turbomachine casing in the region of a combustion chamber, releasing a top half shell from a bottom half shell, axially displacing the top half shell and thereafter removing the top half shell in essentially the radial direction, rotationally fixedly connecting an auxiliary tool to the rotor of the turbomachine, connecting the auxiliary tool to the bottom half shell via axially displaceable adapter slides arranged on said auxiliary tool, axially displacing the bottom half shell to release the bottom half shell from axial connecting contours, rotating the bottom half shell with the rotor by approximately 180° until the bottom half shell assumes essentially the position of the removed top half shell, and releasing the bottom half shell and removal of said bottom half shell.

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 hinged device for at least one door (46.1, 46.2) of an aircraft nacelle (40), whereby the door is able to block an opening (44) that is made in the fairing of the nacelle, the device including a base to which is connected at least one door, able to occupy a first retracted state in which the base is located at least in part inside the nacelle and a second extended state in which the base is located outside the fairing of the nacelle so as to be able to move, in a direction that is essentially parallel to the longitudinal axis of the nacelle, outside of the nacelle by entraining the at least one door (46.1, 46.2) to release the opening (44) at least in part.

Abstract: A turbine engine annular combustion chamber, including longitudinal walls connected by a transverse chamber bottom and a single-piece cowling, the chamber bottom and the cowling each including an inner flange and an outer flange each with a plurality of holes made in it for the passage of fixing systems for fixing the cowling on the chamber bottom. Each fixing system has associated with it at least one tangential slot formed on the corresponding flange of the chamber bottom and/or on the corresponding flange of the cowling, each slot being formed in the immediate vicinity of the corresponding fixing system.

Abstract: The component elements of a jet engine combustion chamber are assembled together in a manner that is designed to facilitate maintenance while also improving aerodynamic performance. In an embodiment, the outer cap and the inner cap have tongues and the chamber end wail has corresponding tongues, and the tongues of the caps are assembled to the tongues of the chamber end wall.

Abstract: A method for fabricating a gas turbine engine combustor that includes a domeplate and at least one fuel injector extending through an opening in the domeplate. The method includes fabricating a heatshield that includes a threaded collar extending upstream from the heatshield, positioning the heatshield on a downstream side of the domeplate such that the threaded collar is received within the domeplate opening, and coupling a retainer to the collar on an upstream side of the domeplate such that the domeplate is securely coupled between the heat shield and the retainer.