Abstract: A fire resistance device intended to be placed between a mounting strut of a double-flow aircraft turbomachine and a connecting cowling equipping this turbomachine, the connecting cowling being intended to connect an upstream ring delimiting an inter-flow compartment, to an arm that extends radially across a secondary flow of the turbomachine. The device includes two contacting lips extending along different lines, of which a first lip with a C-shaped section is integrated within a contact structure, and a second lip including at its end a protruding blocking portion protecting the contact structure.

Abstract: A fire resistance device designed to be placed between a mounting strut of a double-flow aircraft turbomachine and a connecting cowling with which the turbomachine is equipped, the connecting cowling being designed to connect an upstream ring, delimiting an inter-flow compartment, to an arm that extends radially across a secondary flow of the turbomachine. The device can be produced as a single piece and includes two contact lips which extend along different lines, the first lip having a C-shaped cross-section.

Abstract: A hot fairing structure for a pre-diffuser includes a ring-strut-ring structure that comprises a multiple of hollow struts and a multiple of inlets to a respective diffusion passage, one of the multiple of inlets formed between each one of the multiple of hollow struts located between two diffusion passages.

Abstract: Disclosed herein is a seal comprising an annular ring comprising: an outer diameter surface having an outer diameter, an inner diameter surface, a first height measured from the outer diameter surface to the inner diameter surface, wherein a first ratio of the outer diameter to the first height is greater than 10:1; and a cross-sectional shape comprising: a central beam extending from an inner end to an outer end, the central beam including a central axis that defines an angle with a neutral axis of the cross-sectional shape, the angle being between 5 degrees and 25 degrees; a first flange extending axially from the inner end of the central beam in a first axial direction; and a second flange extending axially from the outer end of the central beam in a second axial direction, the second axial direction being opposite the first axial direction.

Abstract: The present application provides an exhaust frame differential cooling system of a gas turbine engine to mitigate a temperature differential along a compressor and/or a turbine to minimize centerline eccentricity of a shaft. The exhaust frame differential cooling system may include a number of compressor temperature sensors positioned about the compressor and/or a number of turbine temperature sensors positioned about the turbine, an exhaust frame including an inner barrel with a bearing tunnel for the shaft, an outer barrel, and a number of struts extending from the inner barrel to the outer barrel, a blower, and a cooling air metering system that provides cooling air from the blower to the bearing tunnel and through the inner barrel, the struts, and the outer barrel in response to the temperature differential being determined along the compressor and/or the turbine.

Abstract: A gas turbine engine and combustor assembly are provided, the combustor assembly including a first liner and a second liner together defining at least in part a combustion chamber, wherein the first liner and the second liner are separated by a gap along the longitudinal direction, and wherein the first liner is forward of the second liner relative to a flow of fluid through the combustion chamber along the longitudinal direction, and wherein the gap is extended along the circumferential direction.

Abstract: A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a seal arc segment including a sealing portion, and a first rail and a second rail opposed to the first rail. The sealing portion extends in a circumferential direction between opposed mate faces and extends in an axial direction between a leading edge and a trailing edge. Each of the first and second rails extend outwardly in a radial direction from the sealing portion to respective first and second edge faces, and the sealing portion has a sealing face dimensioned to bound a gas path and includes a backside face opposed to the sealing face. The first and second rails include respective first and seconds pairs of hooks dimensioned to mount the seal arc segment to an engine static structure in an installed position. The seal arc segment is radially opposed to the sealing face between the first and second edge faces establishing a first region.

Abstract: A vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the flange. There is a conformal thermal insulation blanket disposed on the flange.

Abstract: A fire resistance device is intended to be interposed between an upstream end of an aircraft turbine engine mounting structure and a cowling of the turbine engine delimiting an inter-flow compartment.

Abstract: Fairing assemblies and methods for assembling gas turbine engine fairing assemblies are provided. For example, a fairing assembly comprises a plurality of fairings, an annular inner band defining a plurality of inner pockets, and an annular outer band defining a plurality of outer pockets. Each fairing has an inner end radially spaced apart from an outer end. Each inner pocket is shaped complementary to each fairing inner end and has forward and aft ends. Each outer pocket is shaped complementary to each fairing outer end and has forward and aft ends. The inner and outer bands are each a single piece structure. Each fairing inner end is received within one of the plurality of inner pockets, and each fairing outer end is received within one of the plurality of outer pockets. Some embodiments also comprise an inner ring positioned against the inner band to close the inner pockets.

Abstract: Airfoil assemblies for gas turbine engines are provided. For example, an airfoil assembly comprises an airfoil, an inner band defining an inner opening shaped complementary to an inner end of the airfoil, and an outer band defining an outer opening shaped complementary to an outer end of the airfoil. The airfoil inner end is received with the inner opening, and the airfoil outer end is received within the outer opening. A strut extends radially through an airfoil cavity. A first pad is defined at a first radial location within the cavity. A second pad is defined within the cavity at a second, different radial location. In some embodiments, the airfoil assembly inner band includes a first inner flange, through which the inner band is secured to a support structure, and the outer band includes a first outer flange, through which the outer band is secured to a support structure.

Abstract: A vane arc segment includes an airfoil and a spar that has a spar platform adjacent a first fairing platform of the airfoil fairing and a spar leg that extends through a hollow airfoil section of the airfoil fairing. An end portion of the spar leg has a spar clevis mount. There is a support platform adjacent the second fairing platform that has first and second through-holes. The spar clevis mount protrudes from the support platform. There is a spar pin that extends through the spar clevis mount and locks the support platform to the spar leg. A baffle extends through the spar platform, through the hollow airfoil section, and through the second through-hole of the support platform. The baffle is secured in a joint to the support platform and thereby limits rotation of the support platform about the spar pin.

Abstract: A turbine module includes a turbine rotor assembly, a turbine case, and a turbine shroud assembly. The turbine rotor assembly is mounted to rotate about a central reference axis. The turbine case is spaced radially outward from the turbine rotor assembly circumferentially around the central reference axis. The turbine shroud assembly includes a plurality of turbine shroud segments mounted to the turbine case. Each turbine shroud segment includes a blade track segment that faces the turbine rotor assembly and a carrier mounted to the turbine case.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a spar, a turbine vane, and load transfer pins. The spar comprises metallic materials and is configured to support other components of the turbine vane assembly relative to an associated turbine case. The turbine vane comprises ceramic matrix composite materials and is shaped to include an airfoil configured to direct the flow of hot gasses through a primary gas path of the turbine vane assembly.

Abstract: A blade outer air seal assembly includes a blade outer air seal that has a plurality of segments that extend circumferentially about an axis and are mounted in a support structure. At least two of the segments have a first wall circumferentially spaced from a second wall. A base portion extends from the first wall to the second wall. A first hook extends from the first wall and a second hook extends from the second wall. A wedge seal is arranged between the at least two adjacent seal segments. A clip is configured to bias the wedge seal radially inward.

Abstract: A vane pack for a gas turbine engine includes an annular arrangement of vanes. A ring is secured around the vanes and extends proud of an axial end of the vanes.

Abstract: Provided is a structure for bleeding gas from a chamber for receiving compressed gas from a compressor of a gas turbine engine and supplying the compressed gas to a combustor. The structure includes: a turbine casing covering an outer circumference of a turbine; an engine housing forming, between the engine housing and the turbine casing, a bleeding space that communicates with the chamber; a bleeding duct through which compressed gas in the bleeding space is guided to the outside of the engine housing; and an annular partitioning member arranged on an upstream side, relative to the bleeding duct in the bleeding space so as to separate the bleeding space from the chamber, and having a plurality of communication holes through which an upstream side and a downstream side relative to the partitioning member communicate with each other.

Abstract: A turbine shroud is disposed between stages of stationary turbine nozzles. The turbine shroud includes arcuate shroud segments sealingly engaged with one another. Each shroud segment includes a bottom panel with a radially inner surface and a radially outer surface and opposite ends. Each end includes an end face, an end cutback face spaced from the end face, and an intermediate surface extending between the end face and the end cutback face to define a recess. One or more seals overlap the recess of a circumferentially adjacent pair of shroud segments to produce a sealed joint.

Abstract: Airfoil assemblies for gas turbine engines are provided. For example, an airfoil assembly comprises an airfoil, an inner band defining an inner opening shaped complementary to an inner end of the airfoil, and an outer band defining an outer opening shaped complementary to an outer end of the airfoil. The airfoil inner end is received with the inner opening, and the airfoil outer end is received within the outer opening. A strut extends radially through an airfoil cavity. A first pad is defined at a first radial location within the cavity. A second pad is defined within the cavity at a second, different radial location. In some embodiments, the airfoil assembly inner band includes a first inner flange, through which the inner band is secured to a support structure, and the outer band includes a first outer flange, through which the outer band is secured to a support structure.

Abstract: A turbine assembly for a turbine engine, includes a rotor disk having a plurality of CMC material blades, a stationary nozzle made of CMC material having a plurality of vanes with outer platforms forming a nozzle ring, a turbine ring made as a single piece of CMC material, and a casing of material presenting a coefficient of thermal expansion that is strictly greater than the coefficient of thermal expansion of the ceramic matrix composite material forming the blades of the rotor disk, the nozzle, and the turbine ring, the casing extending around the nozzle ring and the turbine ring, at least the nozzle ring or the turbine ring being connected to the casing via at least one sliding connection having a degree of freedom to move radially.

Abstract: A turbine section for a gas turbine engine includes an independently controllable wheel downstream of the first static vane structure and a turbine rotor downstream of the independently controllable wheel. A method of generating thrust for a gas turbine engine, includes rotating a independently controllable wheel located downstream of a combustor and upstream of a turbine rotor to augment a swirl of a core flow combustion gases.

Abstract: Systems and methods for reducing heat exposure of a turbine casing in a gas turbine engine may be provided. The system may include a blade track coupled with a turbine casing with a clip. The system may further include a nozzle guide vane coupled to the turbine casing. A cavity may be formed by an end of the blade track, the clip, and a portion of the nozzle guide vane. A heat shield may be positioned between the clip and the end of the blade track in the cavity such that an edge of the heat shield and the portion of the nozzle guide vane form a gap. The heat shield and the nozzle guide vane may be positioned such that the gap closes in response to the heat shield and the nozzle guide vane thermally expanding.

Abstract: A turbo-engine module comprises a turbo-engine housing, and a thermal protection shell of the housing. The thermal protection shell is configured to cover at least partially the housing in order to protect same thermally. The module comprises a wear protection strip which is situated between the housing and the thermal protection shell, in a junction region of the housing and the thermal protection shell.

Abstract: A turbine center frame for a gas turbine having a plurality of first components and a plurality of second components alternately mutually adjacently disposed in circumferential direction and bounding a hot gas-conducting flow channel in radial direction, the first components each having two first overlapping portions, and the second components each having two second overlapping portions. In a transition from a first component to a second component, one of the first overlapping portions and one of the second overlapping portions overlap. A centering element, against which, the first components and the second components are braced along the circumferential direction and in the radial direction, centers the first components and the second components relative to a central axis of the turbine center frame.

Abstract: A multi-piece part includes multiple pieces fabricated via different types of fabrication processes, wherein the multiple parts are configured to be coupled to one another to form the assembly. At least one of the multiple parts is fabricated via an additive manufacturing method. The multi-piece part also includes a holder assembly that couples and holds together the multiple pieces of the multi-piece part, wherein the holder assembly comprises a reversible, mechanical-type coupling.

Abstract: An intermediate casing for a turbine engine, comprising at least one central hub and one shroud (12) which are positioned concentrically, and mutually mechanically connected by structural arms (11), with such casing comprising at least one sealing part (13) attached between the shroud (12) and a structural arm (11), with the sealing part (13) including at least one metal core and one seal.

Abstract: An assembly of at least two members. One of the members supports the other member, the assembly defining an axial direction, a radial direction, and a circumferential direction, an inner member of the at least two members being received radially inside an outer member of the at least two members, wherein the inner member and the outer member am attached to each other by a support arrangement, the support arrangement including at least one floating support assembly as a displaceable coupling between an inner member support point provided at the inner member and an outer member support point provided at the outer member. A displacement of support points in a radial direction results in an interrelated relative displacement of the support points in an axial direction end vice versa.

Abstract: A cable conduit may comprise a head end, a sleeve having a foot end, a boot, and a split grommet having a first upper surface, wherein the head end is coupled to the sleeve opposite the foot end, wherein the boot comprises a first flange and is coupled to the foot end, wherein the head end comprises a second flange having a second upper surface and a cutout penetrating into an interior volume of the sleeve, and wherein the split grommet is coupled about the inner diameter of the cutout.

Abstract: A seal arrangement comprising a first and second components made of a ceramic fiber composite material and comprising first and second bounding sections with first and second overlapping sections and first and second securing sections which are connected to and projects from the first and second bounding sections. The first and second overlapping sections are arranged at least partially overlapping such that the first and second bounding sections form an essentially continuous bounding surface. A separation between the first and second securing sections is at least partially sealed by a seal element which comprises at least one first sheet-metal element arranged along the first securing section and at least partially bridging the separation between the first and second securing sections.

Abstract: A steam turbine includes: a stator vane disposed in a cylinder inside a casing through which steam flows from an upstream side toward a downstream side; and a stationary support that supports the stator vane relative to the casing. The stationary support includes: a first supporting body fixed to the casing, a second supporting body that connects the stator vane to the first supporting body; and a replacement body detachably disposed between the first supporting body and the second supporting body on the upstream side.

Abstract: According to an embodiment, a turbine comprises: a cylindrical casing; and turbine stator blades arranged in the casing along a circumferential direction. The turbine stator blades each includes: a blade effective part; a coolant flow path through which a coolant flows in the blade effective part to cool the blade effective part; an outer ring sidewall provided on an outer periphery of the blade effective part; an inner ring sidewall provided on an inner periphery of the blade effective part; and a contact part provided at an end part of the inner ring sidewall with at least part thereof being along a flow direction of a working fluid, coming into contact with the inner ring sidewall of the adjacent blade during operation, and separating from the inner ring sidewall of the adjacent blade when the operation is stopped.

Abstract: A nozzle support system for a nozzle of a gas turbine includes an outer ring, an inner ring disposed within the outer ring, and a plurality of links pivotally attaching the outer ring to the inner ring. Each one of the links may have an outer end that is pivotally attached to the outer ring, and each one of the links may have an inner end that is pivotally attached to the inner ring, such that the inner ring moves along a longitudinal axis of the outer ring.

Abstract: A gas turbine engine and an anti-rotation feature is disclosed. The gas turbine engine includes a fan case, a stator disposed within the fan case, and a wear liner disposed between the fan case and the stator, the wear liner including a wear liner body, and an anti-rotation feature affixed to the wear liner body, wherein the anti-rotation feature includes at least one mating surface to engage the stator to prevent rotation of the wear liner body.

Abstract: A gas turbine engine includes a rotating stage of blades. A circumferential array of blade outer air seal segments are arranged radially outward of the blades. Adjacent blade outer air seal segments provide a circumferential gap. Facing ends of the adjacent blade outer air seal segments include surfaces. A gap seal engages the surfaces and obstructs the circumferential gap. A biasing member is configured to urge the gap seal radially inward toward the surfaces.

Abstract: A retention member for a component of a gas turbine engine and methods of using the same are provided. The retention member includes an annular body having a first side, a second side, a first end, and a second end, a retention element configured at the first end of the annular body and on the first side, the retention element configured to releasably engage with an interior surface of a case of the gas turbine engine, and a support element configured at the second end of the annular body, the support element configured to engage with a surface of at least one of a blade outer air seal or a blade outer air seal support.

Abstract: A turbine exhaust case comprises a frame, a fairing, a heat shield and a mechanical linkage. The frame comprises an outer ring, an inner ring, and a plurality of struts joining the outer ring and the inner ring. The fairing comprising a ring-strut-ring structure disposed within the frame. The heat shield is disposed between the frame and the fairing. The mechanical linkage couples the heat shield to the fairing. In one embodiment, the heat shield comprises a multi-piece heat shield that inhibits heat transfer between the frame and the fairing. In various embodiments, the mechanical linkage comprises a slip joint or a fixed joint for coupling the heat shield to the fairing.

Abstract: A stator vane arrangement for a turbine engine includes a first vane platform, a second vane platform and a plurality of stator vanes extending radially between the first and the second vane platforms. The first and the second vane platforms extend circumferentially around an axis, and the first vane platform includes an aperture. The stator vanes are arranged circumferentially around the axis, and include a first stator vane that extends radially into the aperture and is fastened to the first vane platform.

Abstract: A mid-turbine frame (“MTF”) for a jet engine is disclosed and comprises a duct that extends between a high pressure turbine (“HPT”) and a low pressure turbine (“LPT”), the duct comprising a plurality of segments that together form an outer annular structure and an inner annular structure, the inner annular structure situated radially inward of the outer annular structure, and/or a plurality of vanes that extend radially outward from the inner annular structure toward the outer annular structure, each vane comprising a channel. Each segment may be coupled to an adjacent segment by a seal.

Abstract: A method of damping a vibration in a rotatable member and a damping system for a rotatable machine are provided. The damping system includes one or more damping stages. The rotatable machine further comprising a casing at least partially surrounding the rotor. The casing includes inwardly extending vanes that include a radially outer root, a radially inner distal end, and a stationary body extending therebetween. The one or more damping stages includes a damper supportively coupled between one or more roots of the plurality of vanes and the casing, an air bearing fixedly coupled to one or more distal ends of the plurality of vanes and configured to bear against the rotatable body wherein the damping stage is configured to receive vibratory forces from the rotatable body through the air bearing and the vane and ground the received forces to the casing through the damper.

Abstract: A turbine assembly of an engine is disclosed. The turbine assembly includes a nozzle ring for receiving exhaust gases from a combustion chamber having a plurality of guide vanes for guiding the exhaust gases, a stator ring disposed downstream to the nozzle ring, and a rotor disposed circumferentially within the stator ring. The stator ring includes a plurality of stator vanes adapted to receive the exhaust gases through the guide vanes of the nozzle ring. The stator ring is axially movable between a first position and a second position along at least one cross-key pin disposed between the stator ring and a turbine casing. Each of the plurality of stator vanes moves into a throat plane of each pair of the plurality of guide vanes in the first position of the stator ring, and moves out of the throat plane in the second position of the stator ring.

Abstract: The present invention provides a support duct for a gas turbine engine. The supported duct comprises at least one support frame having an aperture defined by an upper aperture wall, a lower aperture wall and opposing aperture side walls. It further comprises a duct extending through aperture, the duct comprising an upper panel in abutment with the upper aperture wall, a lower panel in abutment with the lower aperture wall and opposing side walls extending between the upper and lower panels. The opposing side walls of the duct are spaced from the opposing aperture side walls.

Abstract: A flow control device for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner surface and a radially outer surface and at least one stand-up protruding from the radially outer surface and configured to seal an interrupted surface.

Abstract: The disclosure relates Frame segment for a transition piece-turbine interface having a picture frame receptacle for axially receiving an aft end of a combustor transition piece. The frame segment can include an I-beam with an upper horizontal element, a lower horizontal element, and a vertical web, wherein the upper horizontal element has mounting face for fixation to a vane carrier. The vertical web has a downstream face, facing towards a first stage of a turbine when installed in a gas turbine. The vertical web includes a cooling gas duct for supplying cooling gas to the downstream face of the vertical web.

Abstract: The present application provides a gas turbine engine. The gas turbine engine may include a compressor discharge casing, a number of combustors configured in an annular array, and a number of aft mounting assemblies. An aft mounting assembly mounts a combustor to an inner diameter of the compressor discharge casing.

Abstract: A turbomachine component is provided having at least two sub-components that are separated by a parting joint and each have a sealing surface at the parting joint, at least one of the two sealing surfaces being convex in order to form a linear contact of the two sealing surfaces. At least one of the sealing surfaces has a coating on it which includes a hard material, is a maximum of 30 ?m thick and is applied using a vapour deposition method, or a coating which includes a chrome-containing alloy, is a maximum of 30 ?m thick and is applied by a vapour deposition method, or is a maximum of 300 ?m thick and applied using a thermal spraying method.

Abstract: A support structure suitable for interposing between the engine and the nacelle of an aeroengine and suitable for being fastened on an intermediate casing is provided. The structure includes a shroud having a plurality of sectors, a plurality of radial arms each installed between two adjacent shroud sectors, and a plurality of fasteners for fastening the plurality of shroud sectors to the plurality of radial arms. The shroud sectors and radial arms form a skeleton for transmitting structural forces within the intermediate casing. The structure includes non-structural fairings mounted on the skeleton and suitable for reconstituting the airflow passage. Each of the shroud sectors and radial arms is formed by a metal frame without mechanical discontinuity, each shroud sector metal frame defining at least one opening that is closed by a cover and each radial arm metal frame defining at least one orifice for receiving a closure plate.

Abstract: In one aspect the present subject matter is directed to a system for thermally isolating a turbine shroud of a turbine shroud assembly. The system includes a shroud support having an inner surface and a turbine shroud that is connected to the shroud support. The turbine shroud includes a hot side surface that is radially spaced from a back side surface. At least a portion of the back side surface is oriented towards the inner surface of the shroud support. The system further includes a coating that is disposed along the back side surface of the turbine shroud. The coating regulates heat transfer from the turbine shroud to the shroud support or other hardware that may surround or be adjacent to the turbine shroud.

Abstract: One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together.

Abstract: Embodiments of a turbine nozzle are provided, as are embodiments of methods for the manufacture of turbine nozzles. In one embodiment, the turbine nozzle includes a support ring and a slip joint ring, which is substantially concentric with the support ring and radially spaced apart therefrom. The slip joint ring has a plurality of slots therein. A plurality of vanes is fixedly coupled to the support ring and extends radially therefrom into the plurality of slots. A plurality of radial slip joints is formed between the plurality of vanes and the plurality slots. Each slip joint extends around a different one of the plurality of vanes to permit relative radial movement between the plurality of vanes and the slip joint ring during operation of the turbine nozzle.

Abstract: A rotor disk for a gas turbine engine is disclosed and formed to enable operation at high rotational speeds in a high temperature environment. The rotor disk is formed to include a bore, a live rim diameter and an outer diameter related to each other according to defined relationships.