Abstract: A combustion chamber of a gas turbine, wherein downstream of the combustion chamber an expansion turbine with a turbine inlet is arranged. The combustion chamber has an annular design with a component, which has a chamber wall and an end wall arranged next to the turbine inlet and a corner as connection of the chamber wall and the end wall. To increase the cooling performance the component further has an air guidance piece arranged at a distance from the chamber wall with a cooling channel in-between. Further, the corner is fluid tight, wherein the distance of the air guidance piece to the end wall is at least 0.5-times and at most 2-times the lowest width of the cooling channel.

Abstract: The invention relates to a straightening assembly (28) comprising two radially inner and outer coaxial shells (34) between which extend vanes (36) made of composite material, fixed to a first end portion on the radially inner shell and to a second end portion on the radially outer shell, characterized in that: for each vane (36), in a plane perpendicular to the axis of the radially inner and outer shells (34), a straight line passing through a junction between said first and second end portions and the useful part forms an angle ? with a radius of the radially inner shell, passing through the junction between said first end portion and the useful part of said vane (36), such that 0°<??30°; the radially inner shell has a diameter ranging from 1,000 mm to 1,600 mm; and in that the radially outer shell has a diameter ranging from 2,000 to 2,800 mm; the number of vanes (36) ranges from twenty-five to forty-five.

Abstract: In one exemplary embodiment, a flow path component assembly includes a support structure that is a unitary component having an axial portion and a radial portion and an outer portion arranged between the support structure and a flow path. The outer portion defines a gap between the outer portion and the support structure axial portion. A flange extends into the gap.

Abstract: A turbine vane assembly includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The turbine vanes are adapted to interact with gases flowing through a gas path of the gas turbine engine. The outer vane support is arranged radially outward of the turbine vane and configured to receive aerodynamic loads from the turbine vanes. The inner vane support is arranged radially inward from the outer vane support to locate the turbine vane assembly therebetween.

Abstract: A system for coupling a shroud to a case associated with a gas turbine engine and a gas turbine engine including such a system includes the case defining a bore and the shroud retained within the case. The shroud defines a pocket. The system includes a pin received through the bore and at least partially positioned within the pocket. The pin has a perimeter. The system includes a load spreader including a first side and a second side opposite the first side. The first side is interconnected to the second side by a flexible portion. The first side, the second side and the flexible portion are received about a portion of the perimeter of the pin, and the load spreader is configured to transmit at least one of an axial point load and a circumferential point load from the pin over a surface of the shroud.

Abstract: A blade outer air seal assembly includes a blade outer air seal segment that has a base portion that extends from a first axial side to a second axial side and from a first circumferential side to a second circumferential side. An intersegment seal is arranged adjacent the first circumferential side. The intersegment seal has a first seal portion and a second seal portion connected by a rib.

Abstract: An airfoil assembly includes a vane that includes an outer platform, an inner platform, and an airfoil. The outer platform defines an outer boundary of a gas path. The inner platform is spaced apart axially from the outer platform relative to an axis and defines an inner boundary of the gas path. The airfoil extends axially between and interconnects the outer platform and the inner platform. Force loads caused by gases interacting with the airfoil are configured to be transmitted from the airfoil assembly to a case arranged around the airfoil assembly.

Abstract: A gas turbine rotor includes: a first rotor in a shaft shape extending by a predetermined length in an axial direction, and provided with an insertion hole at an end of the first rotor; a second rotor in a shaft shape extending by a predetermined length in the axial direction, and provided with an insertion part extending by a predetermined length at an end of the second rotor and corresponding to the insertion hole such that the insertion part is inserted into the insertion hole; and a connecting member mounted between the insertion hole and the insertion part to connect the first rotor and the second rotor together, and made of a material having a thermal expansion coefficient different from a thermal expansion coefficient of both the first rotor and the second rotor.

Abstract: A shroud hanger assembly is provided for hangers and shrouds defining dimensionally incompatible components such as those which are press or frictionally fit to engage one another. The shroud hanger assembly includes a multi-piece hanger and a shroud, which is pinned to the hanger assembly by at least one axially extending pin and which locates the shroud relative to the hanger to control motion in one or both of circumferential (tangential) and radial directions relative to the engine.

Abstract: An airfoil includes an endwall section and an airfoil section that defines, at least in part, an airfoil profile. The airfoil section includes an internal passage and a rib that sub-divides the internal passage. At least one of the rib or the endwall section includes a seal cavity. A seal is disposed in the seal cavity.

Abstract: A gas turbine engine component includes an outer diameter endwall, an inner diameter endwall spaced radially inward of the outer diameter endwall, and at least one body supported between the outer and inner endwalls for rotation about an axis. The body includes an outer diameter surface spaced from the outer diameter endwall by a first gap and an inner diameter surface spaced from the inner diameter endwall by a second gap. The outer and inner diameter surfaces and the outer and inner diameter endwalls are configured such that the first and second gaps remain generally constant in size as the body rotates about the axis.

Abstract: A vane has an airfoil extending between a radially outer platform and a radially inner platform. At least one of the platforms has nominally radially thinner portions, and a pad defining a radially thicker portion. The pad has a radial thickness that is greater than a thickness of the nominal radially thinner portions. The pad surrounds an outer periphery of the airfoil on a side of the radially outer platform. The pad has a varying radial thickness. A mid-turbine frame and a gas turbine engine are also disclosed.

Abstract: Turbine engine turbine sealing module extending about an axis and including a distributor fixed to a casing, and at least one blade connected to an outer platform. The turbine module further includes an impeller mounted rotating inside the casing and surrounded by a sealing ring fastened to this casing. The sealing ring includes an annular row of ring sectors arranged such that the circumferential end edges of two adjacent sectors are facing one another. Each sector includes a body configured to engage with at least one seal lip carried by the impeller and a hook which extends circumferentially which is configured to engage with a fastening rail of the casing. Each ring sector further includes a deflector which extends radially inwards and upstream with respect to the axis, such that the radially inward end thereof extends around a downstream end of the outer platform of the distributor.

Abstract: A gas turbine includes a housing; a rotor rotatable by a fluid flowing through the housing; a bearing for rotatably supporting the rotor; and a support structure configured to support the bearing with respect to the housing. The support structure includes an inner casing accommodating the bearing; an outer casing fastened to the housing; and a strut extending between the inner and outer casings, and at least one of the inner casing and the outer casing includes a diaphragm that is deformable in a radial direction of the rotor. Each casing includes a strut root connected to the strut, a strut platform surrounding the strut root and being formed as the diaphragm; and a main body surrounding the strut platform. Damage to the support structure due to thermal expansion can be prevented, and the support structure can be easily designed to avoid resonance between the support structure and the rotor.

Abstract: A clip for an air seal assembly, comprising a clip base having a control ring interface; a biasing element coupled to the clip base opposite the control ring interface; wherein the clip is insertable into a clip receiver and configured to engage a notch of a blade outer air seal and configured to detachably couple with a clip mounting flange.

Abstract: An apparatus and system for mechanically connecting components are provided. The apparatus includes a mechanical connecting joint that includes a first joint member formed of a material having a first coefficient of thermal expansion (CTE) value, the first joint member comprising a first sidewall, a second opposite sidewall, and a body extending therebetween. The mechanical connecting joint further includes a second joint member formed of a material having a second CTE value, the second CTE being less than the first CTE. The second joint member includes a first leg facing the first sidewall, a second leg facing the second sidewall, and a connecting member extending between the first leg and the second leg. A first gap is formed between the first joint member and the first leg and a second gap is formed between the first joint member and the second leg.

Abstract: Disclosed is a lever linkage for the rotationally fixed connection of a guide vane to a lever of a guide vane adjusting device of a turbomachine, wherein the guide vane has a vane shaft, which extends along a vertical axis. In accordance with the invention, the lever is formed in one piece at a radially outer end of an essentially hollow cylindrical clamping sleeve, which coaxially surrounds the vane shaft in sections, and the vane shaft and the clamping sleeve are coupled by way of a longitudinal side form-fitting connection or a front-end form-fitting connection, and the vane shaft can be tensioned with the clamping sleeve along the vertical axis by means of a fastening element, in particular a threaded nut. In consequence thereof, a separation of two different force flows that act on the guide vanes is obtained, as a result of which local load peaks are reduced.

Abstract: In a gas turbine exhaust member and an exhaust chamber maintenance method, the gas turbine exhaust member is provided with: an inside diffuser that forms a tubular shape and is divided into multiple parts in the circumferential direction; a first seal housing that forms a tubular shape and is integrally formed in the circumferential direction, the front end of which being coupled to the rear end of the inside diffuser; a second seal housing that forms a tubular shape and is integrally formed in the circumferential direction, the front end of which being coupled to the rear end of the first seal housing; and a supporting coupling part that supports the rear end of the first seal housing and the front end of the second seal housing so as to allow the rear end and the front end to move in the axial direction.

Abstract: Turbine vane assemblies incorporating both metallic and ceramic matrix composite materials are provided in the present disclosure. The turbine vane assemblies further include interface components that allow for differing rates of thermal expansion in the ceramic matrix composite components and the metallic components.

Abstract: An actively cooled assembly adapted for use in a gas turbine engine is disclosed herein. In illustrative embodiments, the assembly includes at least one ceramic matrix composite component with an attachment feature.

Abstract: A sealing assembly for a gas turbine is provided, especially an aircraft gas turbine, including: a first component having a first flange portion and a second component having a second flange portion; in the area of the first flange portion and the second flange portion, the first component and the second component being detachably intercoupled by a bolt connection; between the first component, especially a rounded portion, and the second component, especially the second flange portion, a variable interspace being provided that is at least partially sealed by at least one sealing element. The sealing element has a first sealing portion that rests against the first component, especially against the rounded portion, and a second sealing portion that rests against the second component, especially against the second flange portion; the first sealing portion and the second sealing portion being interconnected and extending obliquely relative to each other.

Abstract: A turbine ring assembly includes a ring support structure and a plurality of ring sectors made of ceramic matrix composite material, each ring sector having a portion forming an annular base with an inside face defining the inside face of the turbine ring, and an outside face from which there extends a wall defining an internal housing in which a holder member made of metal material is present, the holder member being connected to the ring support structure and including a body from which elastically deformable holder elements extend inside the internal housing on either side of the body, the holder elements bearing against the wall.

Abstract: A turbine component assembly is disclosed, including a first component, a second component, and an interface shield. The first component is arranged to be disposed adjacent to a hot gas path, and includes a ceramic matrix composite composition. The second component is adjacent to the first component and arranged to be disposed distal from the hot gas path across the first component. The interface shield is disposed on a contact region of the first component, and directly contacts the second component.

Abstract: A sealing structure includes: an insertion hole formed in a sealing plate; a sealing ring provided in the insertion hole; a first rotating body provided to the shaft, at least a part of the first rotating body positioned inside the insertion hole; a second rotating body disposed at a position apart from the compressor impeller than the first rotating body and having a maximum outer diameter larger than a minimum inner diameter of the insertion hole; and an annular receiving groove formed by the first rotating body and the second rotating body and receiving an inner circumferential surface side of the sealing ring.

Abstract: A gas turbine engine component having a cooling passage includes a first wall defining an inlet of the cooling passage, a second wall generally opposite the first wall and defining an outlet of the cooling passage, a metering section extending downstream from the inlet, and a diffusing section extending from the metering section to the outlet. The metering section includes an upstream side and a downstream side generally opposite the upstream side. At least one of the upstream and downstream sides includes a first passage wall and a second passage wall where the first and second passage walls intersect to form a V-shape.

Abstract: A combustor section of a gas turbine engine has a diffuser case with a structural cone having variable wall thicknesses strategically located for reducing localized stress in the cone.

Abstract: An airfoil including an airfoil body, a recessed portion of a first depth in a first side of the airfoil body, the recessed portion including a plurality of pockets of a second depth located within the recessed portion and ribs of the first depth located between the pockets, a cover configured to fit into the recessed portion such that an interior surface of the cover engages the ribs and an exterior surface of the cover is about flush with an exterior surface of the first side of the airfoil body, and a high energy beam weld configuration extending through the cover and into the ribs and positioned to attach the cover to the ribs.

Abstract: A support device supports an inner member of a fixed body including the inner member and an outer member, the inner member being arranged around a rotating body and divided into upper and lower sections and the outer member being arranged in at least a part around the inner member and divided into upper and lower sections.

Abstract: A section of a gas turbine engine includes a rotor blade designed to rotate about an axis. The section also includes a case positioned radially outward from the rotor blade and extending circumferentially about the axis. The section also includes a control ring being annular, positioned radially inward from the case and designed to move radially relative to the case. The section also includes a segmented blade outer air seal (BOAS) including a plurality of BOAS segments each being positioned radially outward from the rotor blade, movably coupled to the control ring, and designed to move circumferentially relative to each other such that a circumferential gap between each of the plurality of BOAS segments changes in size in response to a temperature change in the section of the gas turbine engine.

Abstract: A turbine shroud for a gas turbine engine includes a carrier and a blade track segment. The carrier is formed to include an inwardly opening blade track channel and the blade track segment is positioned in the blade track channel to couple the blade track segment with the carrier.

Abstract: A turbine vane assembly for use in a gas turbine engine includes an endwall, a flow path component, and a load-distribution system. The endwall is arranged around a central axis of the turbine vane assembly. The flow path component is configured to direct fluid flow through the turbine vane assembly. The load-distribution system is positioned between the endwall and the flow path component to distribute loads transmitted between the endwall and the flow path component.

Abstract: A frame employed in a turbine exhaust case of a gas turbine engine includes a frame outer ring, a frame inner ring, and a plurality of frame struts that connect the frame outer ring to the frame inner ring. The frame outer ring includes a plurality of bosses disposed circumferentially around the frame outer ring, wherein at least one of the plurality of bosses includes a mounting surface that receives and secures a removable bracket to the frame outer ring that supports a weight of the frame.

Abstract: The present invention provides a foundation for a wind turbine. To reduce set-up time and to allow complete alignment of the platform which carries the base flange of the tower construction, the invention provides a foundation with a cage structure having an upper stress distribution flange connected by a plurality of tensioned anchor bolts to a lower flange. The flanges are separated by a number of distance elements whereby the shape of the cage structure becomes fixed by the combination between tensioned bolts and distance elements. Since the cage structure has a fixed shape, the upper stress distribution flange can be aligned before the cage structure is embedded in concrete, and it becomes unnecessary to wait for the concrete to harden.

Abstract: An object of the present invention is to decrease a surging limit flow rate at a low flow rate time, by increasing the inflow velocity to the blade of the impeller wheel, by providing a resistive element that narrows in the radial direction a passage cross section of an air intake passage which communicates between a impeller wheel of a centrifugal compressor and an air intake opening. The centrifugal compressor includes: the compressor housing 9 having the air intake opening 13 opened in a rotary shaft direction, and the air intake passage 11; and the impeller wheel 7 for compressing the air flowing in from the air intake opening 13, inside the housing.

Abstract: A turbofan engine structural guide vane is mounted to a forward bulkhead of a core engine case structure at an inner end and to a fan case at an outer end. A plurality of shear pins extend from the aft portion of the structural guide vane into a corresponding plurality of openings defined in the bulkhead for bearing circumferential loads.

Abstract: A seal assembly includes a seal arc segment defining first and second seal supports with a carriage defining first and second support members. The first support member supports the seal arc segment in a first ramped interface, and the second support member supports the seal arc segment in a second ramped interface. A spring is configured to bias the seal arc segment axially. A heat shield is radially inward of the spring.

Abstract: A hollow vane, includes a main portion, of which the cavity is closed by a cover, that is produced by expanding the main portion relative to the cover, for example by means of differential thermal expansion, then by placing the cover at the bottom of the opening of the cavity, before allowing the main portion to close again, closing a nesting between a protrusion at the bottom of the cover and a widening at the bottom of the opening. The disadvantages linked to the use of welding, glue or even assembly parts are as such avoided.

Abstract: A seal assembly includes a seal arc segment that has first and second seal supports. A carriage has first and second support members. The first support member supports the seal arc segment in a first ramped interface and the second support member supports the seal arc segment in a second ramped interface such that the seal arc segment is circumferentially moveable with respect to the carriage. First and second opposed springs bias the seal arc segment toward a circumferential default position.

Abstract: The invention proposes a casing (30) of an aircraft turbine (16), intended to support a set of ring sectors (28) which partly delimits a channel for the passage of a gas stream through the turbine (16), where the casing (30) includes means for dynamic adjustment of the radial position of the ring sectors (28) by controlled injection of an air stream on to portions (36) of an annular wall (34) of the casing (30), where the casing (30) includes an upstream radial tab (38) which connects an upstream end of each ring sector (28), in the flow direction of the gas stream, to the casing (30), and a downstream radial tab (40) which connects a downstream end of each ring sector (28) to the casing (30), where both upstream and downstream radial tabs (38, 40) are made from a single piece with the casing (30), characterized in that each radial tab (38, 40) is made as two portions (42, 44) from different materials.

Abstract: This disclosure relates to a gas turbine engine including a blade outer air seal (BOAS) having at least one attachment hook adjacent one of a leading edge and a trailing edge thereof. The BOAS further includes at least one radial standoff axially aligned with the at least one attachment hook.

Abstract: An inner diffuser case for a gas turbine engine may include a pre-diffuser including a plurality of struts. At least one of the struts may include an aperture therethrough. A cone may have a downstream end coupled to the pre-diffuser downstream of, and proximate to, the aperture. A skirt may have a forward end coupled to the pre-diffuser downstream of, and proximate to, the aperture. The forward end of the skirt may be axially aligned with the downstream end of the cone.

Abstract: A gas turbine sealing assembly includes a first static gas turbine wall, a second static gas turbine wall, and a leaf seal having a first side and a second side. The first static gas turbine wall contacts the second side at a first position, and the second static gas turbine wall contacts the second side at a second position. A spring exerts force on the first side. The spring includes a first spring wall coupled to the first static gas turbine wall. A second spring wall extends radially outward from the first spring wall. A third spring wall extends axially away from the second spring wall. A fourth spring wall extends radially inward from the third spring wall and includes a radially inner end. The radially inner end of the fourth spring wall contacts the first side of the leaf seal between the first position and the second position.

Abstract: A compressor assembly is disclosed. The compressor assembly may have a compressor housing. The compressor assembly may also have a compressor impeller disposed within the compressor housing. Further, the compressor assembly may have a bearing housing attached to the compressor housing. The compressor assembly may also have a diffuser ring disposed between the compressor housing and the bearing housing. In addition, the compressor assembly may have at least one tab. The tab may extend radially outward from the diffuser ring. The tab may be configured to engage with a recess in the compressor housing. The compressor assembly may also have at least one fastener. The fastener may be configured to attach the diffuser ring and the compressor housing.

Abstract: A fixed vane-type turbocharger is provided with a fixed vane in a conduit between a bearing housing and a turbine housing. The fixed vane is configured by a movable member disposed on one of the mutually opposing front faces of the bearing housing and the turbine housing so as to be capable of forward and backward movement, and by vanes which are fixed to the front face of the movable member. A pressing member presses the movable member so that the distal ends of the vanes are brought into pressure contact between the rear face of the movable member and either the bearing housing or the turbine housing with the other opposing front face of these. The pressing member contacts the rear face of the movable member within a range in the radial direction where the vanes are disposed, and presses within this range.

Abstract: An assembled ring shroud is provided. The assembled ring shroud may comprise a cover, a first ring layer, a key layer, and a second ring layer. The cover may operatively wrap around, contain, and secure the first ring layer, the key layer, and the second ring layer. The key layer may comprise a key that may act for position and anti-rotation. The key layer may comprise spacers to hold the keys in place. The first ring layer and the second ring layer may be configured to provide further strength and support to the assembled ring shroud.

Abstract: A gas turbine engine includes a service line interface that extends from an outer surface of a bearing support such that a heat shield is spaced from the outer surface. A method of mounting a heat shield to a gas turbine engine includes trapping a heat shield within a gapped overlap spaced from an outer surface of a bearing support.

Abstract: A method of assembling a gas turbine engine includes locating a consumable assembly tool clip within the engine.

Abstract: A connector between a first object having a first coefficient of thermal expansion and a second object having a second coefficient of thermal expansion has a slot attaching to the second object. The slot formed by an inner shell having a slot shape and a first portion that blends into an outer surface of the second object, the slot having an enclosed portion facing forward and an aft portion that is open.

Abstract: A turbine system is provided having a first turbine casing and a second turbine casing, the first and second turbine casings together defining an inner wall. The turbine system further includes a first attachment flange extending from a surface of the first turbine casing within the inner wall and a second attachment flange extending from a surface of the second turbine casing within the inner wall. The first attachment flange defines a first aperture and the second attachment flange defines a second aperture. A pin extends through the first aperture and into the second aperture.

Abstract: A centerline support bar for steam turbine component assembly, installation, and/or alignment is disclosed. In one embodiment, the support bar includes a body portion configured to be disposed across a centerline of a steam turbine, the body portion configured to transversely connect to a steam turbine component; and a first flange extending from the body portion and configured to contact a steam turbine casing, wherein the steam turbine support bar is configured to non-affixedly join the steam turbine casing to the steam turbine component.