Abstract: A casing assembly for a gas turbine engine includes a plurality of vane casing segments. Each vane casing segment includes an arcuate member and a pair of split-line flanges. Each split-line flange includes a first axial flange end and a second axial flange end. Each split-line flange is fixedly coupled to an adjacent split-line flange of an adjacent vane casing segment. Each vane casing segment includes at least one row of stator vanes welded to the arcuate member. Each split-line flange is at least partially and circumferentially inclined relative to a rotational axis of the gas turbine engine, such that the first axial flange end is circumferentially offset from the second axial-flange end. Further, each split-line flange includes an intersecting portion, such that at least the intersecting portion of each split-line flange is circumferentially inclined relative to the rotational axis.

Abstract: A shroud hanger assembly or shroud assembly is provided for dimensionally incompatible components wherein the assembly includes a multi-piece hanger, for example having a forward hanger portion and a rearward hanger portion. A cavity is formed between the parts; wherein a shroud may be positioned which is formed of a low coefficient of thermal expansion material. The hanger and shroud may be formed of the same material or differing materials in order to better match the thermal growth between the hanger and the shroud. When the shroud is positioned within the hanger opening or cavity, one of the forward and rearward hanger portions may be press fit or otherwise connected into the other of the forward and rearward hanger portion.

Abstract: A rotor, a turbine, and a gas turbine including the same are provided. The rotor includes a pair of disks rotating about an imaginary central axis and arranged parallel to each other in an axial direction, a replaceable self-locking sealing assembly interposed between the pair of disks, and a fastening section disposed on the sealing assembly to fasten the sealing assembly to the disks. The disk includes a sealing slot disposed on an opposite surface to another adjacent disk and a head slot disposed outward from the sealing slot with respect to a radial direction of the to disk. The sealing assembly includes a main body with one end inserted into the sealing slot through the head slot from an outside of the disk and a sealing head disposed on another end of the main body to be seated on an inner wall of the head slot to restrict the main body from being moved.

Abstract: A compressor section includes a compressor casing that has an upper casing portion coupled to a lower casing portion such that a split-line is defined between the upper casing portion and the lower casing portion. A split-line stator vane assembly includes a first split-line stator vane that has a first shank with a first platform portion and a first mounting portion extends radially outward of the first platform portion. A first airfoil extends radially inward of the platform portion. The first platform portion includes a protrusion that extends circumferentially beyond the split-line. A second split-line stator vane having second shank with a second platform portion and a second mounting portion extends radially outward of the second platform portion. A second airfoil extends radially inward of the second platform portion. The second platform portion includes a recess that extends circumferentially away from the split-line.

Abstract: A flexible cartridge assembly can include a flexible shell that includes a flexible portion disposed between a compressor-side portion and a turbine-side portion, where the flexible portion includes a series of arc-shaped cutouts disposed axially along at least a portion of the flexible portion; a compressor-side bearing assembly; and a turbine-side bearing assembly.

Abstract: A component for a gas turbine engine includes a component body formed of ceramic matrix composite lamina and has at least one hook. The at least one hook has an attachment region radially inward of the at least one hook. The attachment region is radially thinner from a hook end of the at least one hook to a remote end, and then becomes radially thicker. A slot is formed through a radial thickness of the at least one hook from the hook end in a remote direction, such that there are two sections of the attachment region. A gas turbine engine is also disclosed.

Abstract: A component for a gas turbine engine includes a mateface with a purge flow interface, the mateface comprises a pocket located in communication with a feather seal slot in a mateface. A vane for a gas turbine engine includes a platform that extends from the airfoil, the platform comprising a mateface with a feather seal slot and a pocket in communication with the feather seal slot, wherein the pocket is of a cross-sectional shape larger than the feather seal slot.

Abstract: A jacket ring assembly for a turbomachine, the jacket ring assembly including a casing part, a jacket ring segment which is adapted to radially outwardly surround a rotor blade ring and to this end is disposed radially inwardly of the casing part, as considered with respect to a longitudinal axis of the turbomachine, and a segmented ring which is circumferentially divided into segments and by which the jacket ring segment is mounted to the casing part, the segmented ring being axially form-fittingly disposed on a form-fitting element of the casing part, for which purpose each of the respective segments of the segmented ring is radially outwardly assembled with the form-fitting element, and the segmented ring forming a supporting seat on which the jacket ring segment is seated and radially inwardly supported with an axially forward end.

Abstract: An airfoil piece includes at least one inner ceramic matrix composite ply which defines an internal cavity of an airfoil section and first and second collar projections. At least one inner ceramic matrix composite ply is continuous through the airfoil section and first and second collar projections, first and second platforms at the first and second ends of the airfoil section. The first and second collar projections extend radially past the first and second platforms, respectively. A vane for a gas turbine engine and a method of assembling a vane are also disclosed.

Abstract: A rotor for a gas turbine having a rotor disk on which there are a plurality of rotor components distributed around the circumference. The rotor disk has a circumferential securing shoulder with a contact face. Retaining faces come to bear against the contact face, each of the retaining faces have a retaining shoulder of the respective rotor component and are designed with a form that complements the contact face. In order to optimize the bearing stresses between the retaining shoulder and the securing shoulder, the retaining face has a smaller radius than the contact face, namely the retaining radius is at least 0.99 times and at most 0.995 times the contact radius. Also provided is an axially extending aperture in the rotor component, the width of which in the circumferential direction is 25% to 75% of the rotor component width in the circumferential direction.

Abstract: A turbine unit for an aircraft turbine engine, comprises a rotor disc (17) axially continued by a rim (36) and carrying rotary blades (18) defining together with the disc (17) channels for air flow (43); an annular flange (33) including a bearing end (34) and defining together with the rim (36) an air passage (42) communicating with the channels for air flow (43); a member (32) for axially retaining the blades applied against the disc (17) by the bearing end (34), this turbine unit being characterised in that at least one element, out of the retaining member (32) and the flange (33), comprises a groove (44) formed facing the other element, out of the retaining member (32) and the flange (33), and into which a sealing joint (45) against which the other element axially bears is inserted.

Abstract: An airfoil assembly includes an airfoil that has an exterior wall that defines an interior cavity. The exterior wall extends between a leading end and a trailing end and an open inboard end and an open outboard end. The exterior wall is formed of a high temperature-resistant material selected from refractory metal-based alloys, ceramic-based material or combinations thereof. A support frame extends in the interior cavity and protrudes from the interior cavity through at least one of the open inboard end and the open outboard end.

Abstract: A method of manufacturing a bearing support structure of a gas turbine engine, includes: obtaining a bearing support and a casing assembly, the casing assembly having first and second casings extending around a central axis and connected together via struts, the bearing support securable to the first casing at attachment points; selecting a distance between the attachment points of the bearing support and the struts as a function of a required stiffness of the bearing support structure; and adjusting a position of the bearing support relative to the casing assembly until the attachment points are distanced from the struts by the selected distance and joining the bearing support to the casing assembly at the attachment points.

Abstract: A boss for a case of a gas turbine engine, has: an elongated body having a proximal end securable to the case, the elongated body extending from the proximal end along a boss axis to a distal end, and a flange secured to the elongated body at the distal end, the flange securable to an accessory.

Abstract: The centrifugal compressor can have a shroud engaged to a case via a plurality of circumferentially interspaced slots and lugs, the slots extending in at least one of a radial direction and an axial direction relative to a rotation axis of the compressor, the lugs slidingly received in a corresponding slot and configured for sliding in the slot in response to thermal growth of the case relative to the shroud.

Abstract: Airfoils for gas turbine engines are described. The airfoils include an airfoil body having a pressure and suction side walls and an internal rib and defining a cavity between, at least, the pressure side wall, the suction side wall, and the internal rib. A baffle insert is arranged within the cavity and includes a sealing portion arranged adjacent to the internal rib of the airfoil body with a gap therebetween. A seal element is arranged within the cavity and located in the gap. The seal element is free to move relative to each of the baffle insert and the airfoil body within the gap. When a pressure differential exists across the seal element, the seal element will sealingly engage between sealing portion of the baffle insert and the internal rib of the airfoil body to block a flow through the gap.

Abstract: An assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a seal arc segment that has a sealing portion and a pair of opposed rails extending outwardly from the sealing portion. The sealing portion includes a sealing face dimensioned to bound a core flow path and has a backside face opposed to the sealing face. The backside face includes a first localized region, a second localized region between the pair of rails, and a third localized region. A support includes a mounting portion and a first interface portion. At least one retention pin is dimensioned to engage the first interface portion of the support and at least one of the pair of rails such that the seal arc segment is carried by the at least one retention pin. A support plate is arranged relative to the support such that the at least one retention pin is trapped between the support plate and the support.

Abstract: Flow path assemblies for gas turbine engines are provided. For example, a flow path assembly comprises an inner wall defining an inner boundary of a flow path and a plurality of pockets therein, and a unitary outer wall defining an outer boundary of the flow path. The unitary outer wall includes combustor and turbine portions that are integrally formed as a single unitary structure. The flow path assembly further comprises a plurality of nozzle airfoils that each have an inner end radially opposite an outer end and define an internal cavity for receipt of a flow of cooling fluid. The inner end of each nozzle airfoil is received in one of the plurality of inner wall pockets and defines an outlet for the flow of cooling fluid to flow from the internal cavity to the pocket, which forms a fluid curtain to discourage fluid leakage from the flow path.

Abstract: A retention assembly for a gas turbine engine is provided. The retention assembly may include a bolt, a spacer, a housing, a fitting component, and a conduit assembly. The housing may define at least one receiving aperture sized to receive a second portion of the shaft. The fitting component may define a first through-hole and a second through-hole sized for a first portion of the shaft and the second portion of the shaft to extend therethrough. The conduit assembly may include a conduit for securing to the fitting component such that the first through-hole is open to the conduit. The bolt, the spacer, the fitting component, and the housing may be arranged with one another to define a captive clearance fit such that a first clearance space is defined between a first surface of the second through-hole and the first portion of the bolt.

Abstract: A turbine shroud assembly for use with a gas turbine engine includes a turbine outer case, a blade track segment, and a carrier assembly. The turbine outer case is arranged around an axis. The blade track segment is configured to define a portion of a gas path of the gas turbine engine. The carrier assembly is coupled to the turbine outer case and supports the blade track segment.

Abstract: A flow channel for a turbomachine, in particular a gas turbine, having a plurality of ribs that are disposed between a radially inner lateral surface and a radially outer lateral surface of the flow channel and are circumferentially distributed; a first rib of the ribs having a first rib thickness and a first rib length, and a second rib of the ribs having a second rib thickness and a second rib length, the second rib length being shorter than the first rib length, and/or the second rib thickness being smaller than the first rib thickness; a spacing in the circumferential direction between the first rib and the second rib adjacent thereto, and a spacing in the circumferential direction between at least two adjacent ribs of the ribs mutually deviating, and at least one of these ribs having a non-deflecting external profile and or an internal structure being disposed therein.

Abstract: A turbine ring assembly includes ring sectors forming a turbine ring and a ring support structure, each sector having, along a cutting plane defined by an axial direction and a radial direction of the ring, a portion forming an annular base with, in the radial direction, an inner face and an outer face from which a first and a second lug protrude, the structure including a shroud from which protrude a first and a second radial clamp from which the first and second lugs are held. The turbine ring includes an annular flange having a first portion bearing against the first lug, and a second portion removably fixed to the first radial clamp, the first portion including radial slits open on a free end of the annular flange and delimiting first-portion sectors.

Abstract: There is provided a turbine casing capable of easily realizing improvement of reliability regarding a connection with an exhaust pipe. A turbine casing according to an embodiment includes an exhaust pipe connecting part formed of ferritic heat resistant steel, to which an exhaust pipe formed of austenitic heat resistant steel is connected. Here, the exhaust pipe connecting part and the exhaust pipe are fastened by using screws.

Abstract: An air seal for a jet turbine engine with an upper stator, lower stator and finned turbine disk. The thermal expansion of the stators may be regulated by a control ring, which has a lower rate of thermal expansion that the stators, to prevent rubbing between the stator and fins.

Abstract: A steam turbine assembling method includes: disposing an upper half partition plate having an upper half partition plate division surface on an inner peripheral side of an upper half casing having an upper half casing division surface to form an upper half assembly; and, after disposing a lower half partition plate having a lower half partition plate division surface capable of abutting against the upper half partition plate division surface on an inner peripheral side of a lower half casing having a lower half casing division surface capable of abutting against the upper half casing division surface, fixing a lower half position defining portion to the lower half partition plate in a state where a lower half abutment surface abuts against the lower half casing division surface and the lower half partition plate division surface to form a lower half assembly.

Abstract: A centrifugal pump, and components thereof, operable at high speeds, is described under the present disclosure. A hard polymer sleeve can be applied to certain surfaces of a seal casing within the pump. If the sleeve is applied along surfaces near the center shaft, then the hard polymer will withstand the forces and pressures of the system. The hard polymer might not be used along the outer diameter, farther from the shaft, because velocities are higher the further out one goes. The current disclosure allows for the use of fluoropolymer in the lining sleeve. The benefits of fluoropolymer have been unavailable in high speed centrifugal pumps because the forces are too great on the periphery of the seal casing. However, the lower speeds along the interior, near the shaft, allow fluoropolymer to be used.

Abstract: An assembly adapted for use in a gas turbine engine has a carrier component and a supported component. The assembly including a mounting system for coupling the supported component to the carrier component and a seal adapted to resist the movement of gasses between the supported component to the carrier component. In an illustrative embodiment, the assembly is a turbine shroud segment for blocking gasses from passing over turbine blades included in the gas turbine engine.

Abstract: Plated polymeric gas turbine engine parts and methods for fabricating lightweight plated polymeric gas turbine engine parts are disclosed. The parts include a polymeric substrate plated with one or more metal layers. The polymeric material of the polymeric substrate may be structurally reinforced with materials that may include carbon, metal, or glass. The polymeric substrate may also include a plurality of layers to form a composite layup structure.

Abstract: A vane assembly for a gas turbine engine is provided having an outer casing and an inner casing radially spaced apart from the outer casing. An airfoil extends radially between the outer casing and the inner casing. The airfoil includes a spar having an outer endwall positioned adjacent the outer casing and an inner endwall positioned adjacent the inner casing. A pressure side and a suction side extend between the inner endwall and the outer endwall. A ceramic matrix composite cover has a pressure side extending along the pressure side of the spar and a suction side extending along the suction side of the spar.

Abstract: A variable-nozzle turbocharger includes a variable-vane mechanism that has an annular nozzle ring supporting an array of rotatable vanes, an insert having a nozzle portion axially spaced from the nozzle ring, and a plurality of spacers connected between the nozzle portion of the insert and the nozzle ring for maintaining an axial spacing between the nozzle portion of the insert and the nozzle ring. The spacers are structured and arranged to mechanically stop the vanes from rotating in one direction past a maximum-open position and to mechanically stop the vanes from rotating in an opposite direction past a minimum-open position of the vanes.

Abstract: An airfoil vane assembly includes an anti-wear device for blocking fretting on the airfoil shell and/or carrier as an interface between the airfoil shell and carrier during relative movement. The anti-wear device can be secured with one of the airfoil shell and carrier while slidingly engaged with the other.

Abstract: An aircraft propulsor that includes a load bearing moveable panel is described herein. In one example, the moveable panel can be coupled to a propulsor structure to receive loads from the propulsor structure. Such loads can include roll and/or torque loads generated by rotation of a core engine of the aircraft propulsor. The moveable panel can be coupled to the propulsor structure through a plurality of coupling portions on one or both of the moveable panel and the propulsor structure.

Abstract: Described is a turbine center frame for a gas turbine, in particular an aircraft gas turbine, the turbine center frame having a radially inner wall and a radially outer wall bounding an annular space through which hot gas flows and each having a contour facing the annular space, the contours describing an inner annular space curve along the inner wall and an outer annular space curve along the outer wall. At least one vane element extends in the radial direction through the annular space and has an axial leading edge and an axial trailing edge, the vane element having an outer axial width.

Abstract: The invention relates to a sealing assembly for a turbine engine, comprising a wheel mounted inside a sectorised ring carrying an abradable material, having at an axial end an annular groove (14) in which a circumferential rail of the casing is engaged, sealing members being partly engaged in a recess (19) of a circumferential edge of a ring sector and another part thereof being engaged in a recess of a circumferential edge circumferentially opposite a circumferentially adjacent ring sector, characterised in that each sealing member comprises an axial end portion (180a) arranged radially between the circumferential rail and two adjacent ring sectors.

Abstract: An improved system, apparatus and method for controlling vane angles in a gas turbine engine, and more specifically, for correcting vane angle error in a gas turbine engine. An active synchronization ring comprises a plurality of micro-actuators coupled to the synchronization ring to correct distortion in the synchronization ring. The micro-actuators apply a bending moment to the synchronization ring to cancel or compensate for synchronization ring distortion. The micro-actuators may be controlled open loop or closed loop. Strain sensors measure ring distortion and provide signals to the controller for closed loop control.

Abstract: A gas turbine engine having a turbine blade tip clearance control system for increasing the efficiency of the engine by reducing the gap between turbine blade tips and radially outward ring segments is disclosed. The turbine blade tip clearance control system may include one or more clearance control bands positioned radially outward of inner surfaces of ring segments and bearing against at least one outer surface of the ring segments to limit radial movement of the ring segments. During operation, the clearance control band limits radial movement of the ring segments, and the turbine blade tips do not have a pinch point during start-up transient conditions. In addition, the smallest gap during turbine engine operation may be found at steady state operation of the gas turbine engine. Thus, the clearance control system can set the gap between turbine blade tips and ring segments to be zero at steady state operation.

Abstract: A method of sizing a cavity in a part and a part made from such method. The method includes forming the part having the cavity, including forming a plurality of protrusions extending within the cavity from at least one internal surface of the cavity, the protrusions having distal ends bordering an unobstructed portion of the cavity, the unobstructed portion having an initial dimension at least partially defined by a position of the distal ends, pressing a deforming element against the distal ends of the protrusions to plastically deform the protrusions toward the at least one internal surface of the cavity and increase the initial dimension to a final dimension, and disengaging the deforming element from the distal ends.

Abstract: A gas turbine engine combustion chamber includes upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. A frame structure at the downstream end of each segment has multiple spaced radially extending holes. The downstream end of each segment has an axially upstream extending groove and the downstream ring structure has an annular axially upstream extending hook which locates in the groove of each segment. A portion of the downstream ring structure abuts the frame structure of each segment. The downstream ring structure has multiple holes through the portion abutting the frame structure and segment is removably secured to the downstream ring structure by multiple fasteners locating in the holes in the segments and the downstream ring structure.

Abstract: An annular finger seal includes a base portion defining a base edge of the annular finger seal and having a center pass-through aperture. The base edge may be substantially continuous. The annular finger seal may also include a finger portion defining a finger edge of the annular finger seal and including a plurality of fingers. The finger edge may be segmented by gaps disposed between adjacent fingers of the plurality of fingers. A gas turbine engine includes a combustor section, a turbine section, and an annular finger seal. The annular finger seal includes a plurality of fingers and the annular finger seal is disposed at an adjoining interface between the combustor section and the turbine section. The annular finger seal may be mounted to the turbine section and may engage a radially inward facing surface, relative to the engine central longitudinal axis, of the combustor section.

Abstract: An assembly for gas turbine, includes a first part and a second part installed circumferentially around a longitudinal axis of the turbine, where the first part has a first side face adjacent to a second side face of the second part, and where a first aperture made in first side face is facing a second aperture made in second side face, a sealing plate fitted inside the first aperture and the second aperture, at least one longitudinal face of the sealing plate includes at least two projections made either side of a longitudinal central area of the sealing plate, wherein each projection is of a height chosen so as to reduce a clearance between the sealing plate and the first and second apertures.

Abstract: The invention relates to a guide vane segment for aircraft gas turbine, which extend along a respective circular arc and together form an annular section, wherein, in the radial direction between the outer shroud and the inner shroud, a plurality of guide vanes are arranged adjacently in the peripheral direction, the outer shroud and the two sealing walls form a trough-like profile in longitudinal section, wherein, at the axially front or/and rear sealing wall element(s), at least one relief gap with a main section is provided, which extends substantially radially inward, starting from a radial outer edge of the respective front or/and rear sealing wall element(s) along the sealing wall element. According to the invention, it is proposed that the relief gap has at least one additional section that adjoins the main section radially inward, wherein the additional section is formed by at least one curved subsection.

Abstract: An axial compressor comprises a plurality of compressor stages positioned axially adjacent one another within a casing. Each of the plurality of compressor stages comprise a rotor segment and a banded stator segment positioned axially adjacent the rotor segment. An annulus is formed between the casing and an outer flowpath ring of the stator segment. A vane of each stator segment comprises a member extending into the annulus.

Abstract: The invention relates to a guide vane intended to be mounted in a turbine engine between an inner shroud (17) and an outer shroud (16), comprising a longitudinal straightening body (41) for an air flow extending between a first end (42) intended to be positioned at the inner shroud (17) and a second end (44) intended to be positioned at the outer shroud (16), the longitudinal straightening body having an aerodynamic profile defined by a leading edge (41a) and a trailing edge (41b) in the flow direction of the air flow, and by a camber line (41c) extending from the leading edge (41a) to the trailing edge (41b).

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a first turbine rotor comprising an inner shroud, an outer shroud outward of the inner shroud in the radial direction, at least one connecting airfoil coupling the inner shroud and the outer shroud at least partially along the radial direction, and an outer band outward of the outer shroud in the radial direction and extended at least partially in the circumferential direction, and a plurality of connecting members couples the outer shroud and the outer band.

Abstract: A pressure wave supercharger for compressing fresh air for an internal combustion engine includes a cold gas housing, a hot gas housing, and a rotor casing inside which a rotatable cell rotor is disposed. The hot gas housing has a high-pressure exhaust gas duct and a low-pressure exhaust gas duct, while the cold gas housing has a fresh air duct and a charge air duct. The high-pressure exhaust gas duct, the low-pressure exhaust gas duct, the fresh air duct and the charge air duct are fluidically connected to the cell rotor. The hot gas housing has a heat exchanger which is designed in such a way that at least a first bearing for a rotor shaft can be cooled.

Abstract: A structure in a gas turbine engine comprises a spar and a CMC component adjoining the spar and separated from the spar by a cavity supplied by cooling air. At least one rope seal is installed in the cavity within a groove made in the spar to thus compartmentalize the cavity and control the flow of cooling air.

Abstract: Embodiments of the invention relate to a turbine assembly including a shroud assembly and a nozzle assembly axially adjacent to each other which both include confronting radial sides and confronting axial ends. A spline connector having a circumferential portion and an axial portion such that the circumferential portion of the spline connector extends across the axial ends and the axial portion of the spline connector extends across the confronting radial sides.

Abstract: A turbine shroud for incorporation in a turbine of a gas turbine engine has a plurality of butted shroud segments circumferentially arrayed to form a ring. Each of the shroud segments has an arcuate main shroud body portion, a radially inward extending annular flange attached to a first end of the main body portion, and a radially outward extending flange with a plurality of mounting apertures attached to a second end of the main shroud body portion. A first mounting aperture is sized smaller than an adjacent second mounting aperture in the radially outward extending flange.

Abstract: The invention concerns a support casing (24) for a rotor (12) of a turbine engine such as a ducted fan turbojet engine used for propulsion of an aircraft. The casing (24) comprises: an outer annular wall (38) with an inner annular surface (44); an inner hub (40) able to support the rotor (12) of the axial turbine engine and comprising an outer annular surface (42); an annular passage (46) between the annular wall (38) and the inner hub (40); an annular row of arms (48) passing radially through the annular passage (46). Each arm (48) of the casing (24) comprises an orifice (50) arranged in the annular passage (46) radially at the level of one of said annular surfaces (42; 44). Inserts are fitted to the orifices (50) to control the flow passing through.

Abstract: A flowpath apparatus for a gas turbine engine includes: a plurality of ducts arranged in an array, each duct including a peripheral wall structure having a closed perimeter that defines a flow channel from an upstream end to a downstream end thereof; and a support structure positioning a the plurality of ducts in an array configuration.